“Calibrating” the Ballpark Estimating Method

As stated elsewhere,  the methods I have been using are only ballpark bounding calculations.  To “calibrate” how “good” these techniques really are,  at least for launch to low Earth orbit (LEO),  I ran a generic two-stage rocket broadly similar to the Spacex Falcon-9,  before they upgraded it.  This older version of Falcon-9 was rated at 13 metric tons to LEO,  at low inclination out of Cape Canaveral.  Falcon-9 was a little heavier than 500 metric tons at ignition. 

The key assumptions were 5% gravity loss,  and 5% drag loss for the first stage,  and 5% gravity loss only for the second stage,  on a fast ascent trajectory,  plus 4.5% stage inert weights,  stage payload included in that accounting.  I assumed stage 1 burnout outside the “sensible atmosphere”,  at 3.05 km/s achieved velocity.  LEO was assumed to be 7.79 km/s achieved velocity.  I completely ignored the “boost” effect of the Earth’s eastward rotation.

I did some crude engine ballistics based on characteristic velocity c* at 1000 psia chamber pressures (c* = 5900 ft/s for LOX-RP1,  from the vintage 1970 version of the old Pratt & Whitney “Vest-Pocket Handbook”),  and a bell divergence-corrected thrust coefficient CF chart computed for specific heat ratio 1.20.  The divergence thrust correction factor is 0.983,  pretty much an average 15-degree half angle,  and equivalent to most modern curved expansion bells. 

The first stage is expanded “perfectly” to 14.7 psia (101.3 Kpa,  1013 mbar,  760 mm Hg) backpressure.  The chart says CF = 1.57 at expansion ratio 9.00 when read at pressure ratio 68.  There is a simple relationship among CF,  c*,  and Isp,  leading to Isp = 287.9 s for the first stage.  A very slightly-oversimplified model then estimates exhaust velocity as 2.835 km/s. 

For the second stage,  the ambient backpressure is zero (out in the vacuum).  I looked at the chart for pressure ratio set to an arbitrary 1000:1,  and got CF = 1.826 at expansion ratio 65:1.  For that same 1000 psia LOX-RP1,  c* = 5900 ft/s,  I got Isp = 334.8 s,  and exhaust velocity 3.284 km/s. 

5% gravity loss plus 5% drag loss is a total 10% loss for the first stage,  making the effective required velocity not 3.05 but 3.355 km/s.  That corresponds to a mass ratio of 3.2820,  and a propellant fraction of 0.6953.  The corresponding stage payload fraction is 0.2597.  For a nominal liftoff weight of 500 metric tons,  the stage 1 payload (stage 2 plus the “real payload”) is 129.85 tons.  That ignores any interstage weights,  which might be around a ton,  for a second stage ignition weight of 128.85 tons. 

5% gravity loss on the second stage velocity increment of 4.74 km/s results in an effective velocity requirement on the second stage of about 4.977 km/s.  At the higher second stage Isp,  the mass ratio is 4.55183,  and the propellant fraction for that stage is 0.7803.  For the same stage 4.5% inert fraction,  that leaves 0.1747 for the stage 2 “real payload” fraction.   That’s near 22.51 metric tons for a stage 2 ignition weight of 128.85 tons. 

That payload has to ride inside some sort of protective,  aerodynamic shroud.  For the sake of argument,  assume that shroud also weighs about 1 ton.  Therefore,  the real delivered payload is nearer 21 metric tons for a vehicle massing 500 tons at launch. 

The real Falcon-9 is a little over 500 tons at ignition,  and is rated to deliver 13 metric tons to LEO.  My 21 tons is in the ballpark,  clearly,  but still quite a ways “off” for purposes of “exact” estimates.  All in all,  I’d say these ballpark estimating techniques are actually quite good,  especially for relative-comparison calculations.  But it takes a better model than this,  to really “pin things down”. 

For example,  use the 5%-5% loss factors on the entire trajectory to LEO,  with 4.5% inerts,  and the lower-performing first stage engine performance,  but as a single-stage to orbit vehicle.  It really does calculate as technically feasible,  just at only 1.48 ton payload,  and with propellant tanks “stretched” by 37% (volume) to maintain a 500 ton ignition weight. 

That 1.48 tons gets compared to the 21 tons for the 500 ton two-stage bird.  For one stage versus two,  the launch cost could only be factor 2 lower for the one-stage bird,  at the very most.  On a per-unit-delivered payload basis,  the one stage version will always deliver less mass for a higher cost.  That basic effect is why staging was invented,  over 6 decades ago. 

Government Shutdown, Default! Again? NO!!!

A good friend asked me a question regarding near-term Republican election outcomes in light of the current government shutdown.  I responded to him,  and that is reproduced here,  with very little editing,  except for improved clarity. 

Bear in mind that my response is colored by being very angry with both parties,  just more so the Republicans in recent years.  Myself,  I am a die-hard independent,  so do NOT try to stereotype me politically!

My friend’s question (unedited):

“Do you think that the GOP will lose about 90% of their seats after this?”

My response (edits-for-clarity in red):

I wish,  but I don't think so.  There are too many ignoramuses around us who have been brainwashed into believing all the lying propaganda.  Classic Nazi and Bolshevik "big lie" technique.  Weak minds and all that. 

What's really going on is that our two-party system has been transformed into a two-and-a-half party system.  Those so-called "tea party" idiots are an extremist sub-set of the GOP who could not possibly get the votes to pass any of their agenda.  They are a very loud but rather small minority,  and they know it. 

Over the last several years,  they've effectively purged many non-tea-party GOP members from the party,  until they are somewhere close to half its membership.  Once they (the tea party extremists) comprised about half of the surviving GOP,  they could effectively blackmail the non-tea-party types into voting their way,  by mounting primary challenges when they (the non-tea party GOP) didn't vote for the tea party extremist agenda. 

Blackmail and extortion are crimes,  by the way,  in all other walks of life,  except apparently not in politics.  Maybe they should be. 

So,  with control of the House after that massive propaganda campaign leading up to the 2010 election,  they (the tea party extremists) can now force the entire GOP membership of the House to vote in a way that tries to extort the Senate into passing the tea party agenda,  by threatening either government shutdowns or defaults on government debt. 

The lesson of the 2012 election is completely meaningless to them (the tea party extremists),  because they now have a tool by which to extort their way.  Ted Cruz is the leading tea party candidate for 2016 now,  but he's a fake,  just playing the photo ops to get well-known nationally.

Cruz is not there to do the people's business at all,  he just wants to run for President.  That's deliberate,  intentional dereliction of duty,  and I think that is a crime,  too.  Or it ought to be. 

Considering that the Constitution mandates that Congress (House and Senate together) fund the government and pay its bills,  as part of their assigned duties,  it seems to me that threatening not to do its duty is a Constitutional crime being committed by the tea party-controlled portion of Congress.

And they didn't invent it.  Back in 1996,  Newt Gingrich and his "radical Republicans" invented using extortion in Congress,  with that government shutdown back then.  Leading that extortion back then is why I have never liked Newt Gingrich,  then or ever since. 

This tactic (a minority extorting its way) is a crime against the people. 

I'd just as soon see Obama send in the federal marshals and arrest the perpetrators of this travesty,  on charges of wantonly and deliberately failing to do their Constitutionally-mandated duties.  I might include the rest of the GOP for caving-in,  and not resisting this evil.  And I’m really that angry about it. 

Any time party agenda trumps the national interest,  we have the wrong people in office.  Period.  Applies to both parties,  but of late the GOP is primarily the culprit,  and very egregiously so. 

My grandfather told me back in 1965 before he died that the tools for good government are tar,  feathers,  guns,  and ropes.  He said that in his younger days,  folks still used them,  and they had fairly good government.  He also said that in my dad's younger days,  we let the judges and lawyers talk us out of using those tools.  And now look what we have!  (Really bad government.)

Maybe that (not using the tools for good government) ought to change,  too.

Guess why there's such a push for gun control?  We gotta have 'em,  in order to successfully use the tar,  the feathers,  and the ropes!  That push for gun control is what I despise most about the Democrats. 

They're (politicians of all types and parties) still at least a little afraid of us,  even after over half a century of no one policing their behavior.  They need to be very substantially afraid of us,  of what we would do to them,  if they don't do the people's business properly.  And we need to do it to them,  too.  Make a public example of one,  and the rest fall into line,  for a while. 

These "town hall meetings" that a few reps still hold are nowhere near as confrontational as they should be.  The rest of the reps are afraid to hold them.  Have you noticed that?

My friend responded thusly (spelling corrected):

“Obama should sign an E.O. stopping all congressional pay and benefits also.”

My sentiment:

Well,  that would be a start,  but nowhere near an adequate response. 

Wake up,  folks!  This crap has to stop!

GW

Single Stage Launch Trade Studies

Several friends and on-line correspondents (all of whom are also interested in space flight) have been discussing how to launch single stage to Earth orbit (SSTO),  with an eye toward reusability.  Both traditional launch rockets and winged launchers that are effectively spaceplanes are investigated here. 

All items considered here are vertically-launched.  The depressed trajectories flown by horizontally-launched designs are completely different,  and cannot be analyzed in this way at all.  See Figure 1 (all figures at end) for the assumed trajectory shape,  and associated bounding analysis assumptions.

I personally think reusability will “cost” the extra weight to make the structure robust enough to fly multiple times.  Otherwise,  from a rocket propulsion standpoint,  the typical rocket performance levels available to us are:

LOX-RP1……………………305-310 sec Isp

LOX-liquid-CH4…………near 350 sec Isp

LOX-LH2……………………near 460 sec Isp

1972-vintage NERVA…near 900 sec Isp

Note that NERVA is a solid core nuclear device.

I did this as a parametric bounding analysis,  based on the simple rocket equation and some convenient simplifications to support it.  The supporting calculations are summarized in Figure 2.  I looked at inert structural fractions from 5% to 40% (in increments of 5%) as the independent variable,  with required Isp as the dependent variable.  The parameter was payload fractions from 2 to 10%,  in increments of 4%. 

I did not look at ramjet-assist or any other type of airbreather-assisted vertical launch.  The analysis required to support usable trade studies with airbreathers goes well beyond this kind of rocket equation-based bounding analysis. 

The basic results are presented in Figure 3 as parametric curves.  Required Isp (as the ordinate) to accomplish a launch mission that effectively requires 8.56 km/s delta-vee,  is plotted versus inert structural fraction as the abscissa.  The parameter is imposed dead-head “payload” fraction from 2% to 10% by 4% increments. 

Linear interpolation between payload fractions is clearly permissible.  Horizontal lines have been added to represent the available Isp levels,  as described above for the four “realistic” types of rocket propulsion listed above. 

“Dead head” payload includes the real delivered payload,  plus any shroud or capsule weight,  as appropriate.  The structural inert weight includes basic tankage or airframe structures,  plus engines,  plus any recovery equipment or propellants that might be required. 

One-Shot One-Stage Rockets

For one-shot single-stage rocket boosters,  the inert structural fractions can resemble those of currently-flying vehicles,  two-stage or otherwise.  Those range from 5 to 10%,  and are probably closer to 5% in new designs today,  at least with dense propellants that are not “extreme” cryogenics.  With that range of inerts “spotted” on the graph,  Figure 4 gives the trade study results. 

The LOX-LH2 propellant choice gives the “best” results throughout the 5-10% inerts range,  provided that the propellant volumes can be reconciled with a 5% drag loss.  One might “guess” about a 7% dead-head payload allowance,  that might reconcile fairly well with perhaps 7% inerts (very voluminous LH2 tankage with extra insulation). 

Vehicles like that can carry large payload masses inside a fairly-lightweight shroud (say near 1% of launch weight,  leaving the remainder of the “dead head payload” fraction as real payload delivered to orbit).  Or,  they might carry smaller payloads inside one-way or returnable capsules,  such as Orbital’s Cygnus or Spacex’s Dragon,  respectively.  For the sake of argument,  assume 80% of a one-way capsule’s weight might be real payload,  and 60% of a returnable capsule’s weight might be real payload. 

For that 7% deadhead payload LOX-LH2 sizeout,  if shrouded,  then about 6% of the launch weight might be real delivered payload.  If instead a one-way capsule,  then again about 5.6% of the launch weight might be real delivered payload.  If instead a returnable capsule,  then about 4.2% of the launch weight might be real delivered payload. 

LOX-RP1 is just barely infeasible as shown in Figure 4,  but 5% inerts and 4% dead-head payload is feasible with LOX-CH4.  If shrouded,  then perhaps 3% of the launch weight might be real deliverable payload.  If a one-way capsule,  then about 3.2% of the launch weight might be real deliverable payload.  If a returnable capsule,  then about 2.4% of the launch weight might be real deliverable payload. 

For something comparable to a Falcon-9,  the launch weight would be in the neighborhood of 500 metric tons.  The launch price would be near $56.5M.  (Falcon-9 lists as $4300/delivered kg.)  Using these values and the percentages in the preceding paragraphs,  I get:

Propellant…………….inerts%............deadhead%

LOX-LH2……………….7…………………….7

Type…………………….deliv%...............$/del.kg

Shroud…………………6……………………..1880

1-way capsule………5.6………………….2020

Returnable cap…….4.2………………….2690

 

Propellant…………….inerts%.............deadhead%

LOX-CH4……………….5…………………….4

Type……………………..deliv%..............$/del.kg

Shroud………………….4…………………….3770

1-way capsule……….3.2…………………3530

Returnable cap……..2.4………………….4710

Any of the LOX-LH2 configurations would then seem to offer slight cost advantages per unit delivered payload,  over the LOX-RP1 two-stage-to-orbit (TSTO) baseline Falcon-9.  (This baseline is based on Spacex website data as of 9-24-13.)  The LOX-CH4 data are less advantageous than LOX-LH2,  because of the lower Isp performance.  The shroud and 1-way capsule versions seem to offer very slight advantages over the Falcon-9 baseline,  but the returnable capsule seems to be a little less cost-effective. 

Really,  at this level of analysis,  all the LOX-CH4 data are effectively the same unit price as baseline,  and the LOX-LH2 only very slightly better than baseline.  This looks attractive only for a clean-sheet-of-paper LOX-LH2 design.  Otherwise,  the LOX-RP1 TSTO baseline that we have is better. 

Re-Usable One-Stage Rockets

This is a “screwy” case.  It all boils down to what one believes that the realistic effective inert weight fractions might be.  The trade study results are given in Figure 5,  on which I have spotted the roughly 10% inert fraction of Space Shuttle SRB’s,  which are 900-psi pressure vessels,  being solid motor cases,  yet of limited demonstrated reusability. 

My own guess for the inert fractions of fully-reusable liquid stages is closer to the 15-25% range also spotted on the figure.  This “budget” includes not just the tankage and engines,  but also all the necessary recovery equipment (such as chutes and landing legs),  plus a considerable amount of retro-thrust propellants (if a powered descent is the approach taken,  as in Spacex’s “Grasshopper”).  

This might actually be a “low-ball” estimate,  since entry is so demanding an environment.  But it doesn’t really matter.  The curves show basic infeasibility for all three chemical rocket choices,  with the possible exception of LOX-LH2 at only 1% “dead-head” payload.  Such a payload would have to ride the booster “naked”,  as there is no allowance available for a shroud.  No capsule options seem feasible. 

That leaves you only with the nuclear rocket option “NERVA”,  which at 20% inerts could probably carry 13-14% dead-head” payload.  Actually,  considering the relatively low engine thrust/weight for NERVA-type engines,  we’d be lucky to obtain 35% inerts at 2% dead-head” payload.  That would be about a 1% real delivered payload fraction,  inside a shroud,  as the only feasible option.  That’s 5 metric tons delivered,  at the “same $56.5M” price,  for about $11,000/delivered kg.  That’s not very attractive. 

But,  in any event,  to be re-usable means you are flying back to Earth an already-fired nuclear reactor engine,  and you are doing this multiple times.   There are some very serious safety concerns with such an approach.  I really don’t recommend this for Earth surface launch. 

The bottom line is that a re-usable SSTO booster is technically attainable with nuclear rocket propulsion,  but nobody will like the safety risks.  I did not look at re-usable first stages for a chemical TSTO system.  That is what Spacex is really looking at. 

Re-Usable One-Stage Rocket Spaceplanes

Winged rocket spaceplanes that launch by vertical takeoff (VTO) as SSTO,  but return to horizontal landing (HL) have been a longstanding dream.  Again,  the driving assumption is what you believe a realistic inert weight fraction might be. 

Being a winged airframe,  this is the vehicle that most closely resembles an airplane as we have known them for over a century.  Most modern transports and bombers fall in the 40-50% inert weight fraction range,  with carrier-capable Navy “birds” pushing 60% inerts.  That would be for traditional metal construction.  Airframes like that are usually designed for 40,000+ landings and takeoffs. 

You cannot replace all of the metal structures with composite materials.  These are very intolerant of heat.  Not only orbital descent,  but also ascent,  are rather vicious aeroheating environments.  But,  the number of landings and takeoffs might be in the 100-1000 range,  which eases somewhat the robustness (and inert weight) required of the design. 

A “reasonable guess” might be half composites and half metallic,  for a minimum-credible reusable inert weight fraction in the range of 25 to 30%.  Accordingly,  I showed inert fractions from 25 to 40% on the trade study results given in Figure 6. 

All the chemical options are quite clearly infeasible.  Only a nuclear spaceplane powered by some version of a NERVA (or better) would be feasible.  This brings up (again) all the safety concerns of flying back to Earth with a fired nuclear reactor core,  as discussed above for reusable rocket stage boosters. 

Allowing for the low engine thrust/weight ratio of NERVA,  we might achieve 35% inerts at 2% payload fraction.  No shroud or capsule is required,  so the delivered payload is 2%.  That’s 10 metric tons for a 500 ton launch weight.  Again,  assume the same launch cost of $56.5M for the 500 ton nuke,  and you get around $5700/delivered kg.  It does not seem to offer any cost advantage over what we are doing right now:  the Falcon-9 one-shot TSTO calculates as $4300/delivered kg. 

Options Not Considered Here

I have not looked at airbreather-assist for VTO SSTO,  or any depressed-trajectory SSTO and TSTO systems (whether airbreather-assisted or not).  (I have actually looked at the latter,  but not in a way that I trust yet.)  The airbreathers,  particularly ramjet,  require substantially-more sophisticated performance-estimation methods than the simple rocket-equation stuff presented here.  Those are destined for a future article.

Final Comments

One-shot VTO SSTO rocket-stage systems seem to be marginally attractive (relative to a one-shot LOX-RP1 VTO TSTO baseline) from a delivered payload unit cost standpoint,  but only if a LOX-LH2 system is considered in a clean-sheet-of-paper design.  LOX-CH4 seems to offer no real improvement,  and one-shot LOX-RP1 VTO SSTO seems to be essentially technologically infeasible.

Re-usable VTO SSTO rocket-stage systems appear to be completely infeasible for all known chemical propulsion choices,  relative to the one-shot LOX-RP1 VTO TSTO baseline.  A NERVA-type nuclear approach appears to be technically feasible,  but at lower payload fraction due to the low engine thrust/weight inherent with solid core nuclear engines.  Assuming the same basic launch cost for the same launch weight class,  the unit price for delivered payload appears to be more expensive,  relative to the one-shot VTO TSTO LOX-RP1 baseline. 

 

For VTO SSTO rocket spaceplanes,  only the NERVA (or better) option looks to be technically feasible.  Under the same price/launch weight assumptions,  the unit price for delivered payload looks at-best more-or-less comparable to the one-shot LOX-RP1 VTO TSTO baseline,  probably more expensive.

 

Figure 1 – Basic Trajectory and Assumptions

Figure 2 – Basic Calculations and Related Conditions

 

 

Figure 3 – Basic Parametric Rocket Equation Results

Figure 4 – Basic Results for One-Shot One-Stage Rocket Launchers

 

Figure 5 --  Basic Results Revisited for Re-Usable One-Stage Rocket Boosters

 

Figure 6 – Basic Results Revisited for Re-Usable One-Stage Rocket Spaceplanes

Update 9-29-13:

For those not so familiar with rocket work,  these plots can be a little confusing or misleading.  First:  these are for single-stage operations only.  You cannot use these directly for staged vehicles.  Nor can you do anything useful with these plots toward airbreathing-assist,  it's just too coarse for that,  although concepts can be illustrated. 

For airbreathing-assist,  you have to "account" for highly-variable airbreather Isp effects,  how much of the thrust is airbreather,  and what fraction of the whole trajectory is actually assisted by the airbreather.  You also have to worry about having enough thrust to take off,  and that these charts embody only vertical takeoff on a fast ascent trajectory. 

Second,  the slanted curves are just physics as embodied by the classic rocket equation.  There's only 3 categories of vehicle mass considered here:  inerts,  propellant,  and dead-head payload.  The curves show the interplay among the three,  with two explicitly shown,  calculated to a fixed velocity-change requirement.  None of those curves would ever change,  given the same velocity requirement.

The horizontal lines represent the performance levels of typical rocket propulsion technologies.  In essence,  this is the influence of that portion of the mass budget that is propellant.  I showed 3 chemical and one old nuclear system as a guide. 

Technologies can improve,  shifting these horizontal lines slightly,  but chemistry has been "stalled" for decades,  pretty much where it is depicted.  The nuclear technology offers the most hope of improvement,  but has not been seriously worked-on in 4 decades.    What I show is what was cancelled right before it could be flight-tested,  the variant that was most mature back then. 

The vertical lines represent the effects of materials and construction techniques upon the inert weight.  This has seen the most change in recent decades.  The modern 5-10% inert range is now pretty typical of commercial launcher stages.  Rolled textured aluminum alloy panels are what make this possible,  in concert with higher-tech versions of the engines that have lower engine weight for the same thrust.  Long ago,  that was closer to 20% with things more like frame-and-stringer type construction. 

I have to caution readers and users of these graphs that these 5-10% inert weight percentages are typical of one-shot (throwaway) stages,  not anything that might be reusable.  One-shot designs contend with ascent loads and ascent heating only.  Descent loads and descent heating are not only worse,  they are totally different in character.  You have to deliberately design for them from the outset in a reusable design.  You also have to have a service lifetime in mind for a reusable design,  something totally different than "just-surviving-the-mission" with a one-shot design. 

The early history of aircraft design is the most recent example of a technology arena where we have learned a very fundamental lesson the hard way (with many lives lost):  the robustness of a long service life is simply heavier,  because more materials are required to withstand the forces.   There is no escaping that fact-of-life,  and that is why I spotted recent modern aircraft values on the figure 6.  These are basically dry weight divided by max gross weight.  The difference is really both payload and fuel together.  (Airplanes are different from rockets,  after all.)

The 50% I show as "typical" of a long-life transport or bomber aircraft might not be representative of a reusable winged space launcher,  but the 40% of the all-metal X-15 rocket airplane is a good startpoint for guessing what might be suitable for a reusable winged craft.  Those are fundamentally different from "not-winged" vertical launch stages,  reusable or not.   

Composites typically have at least twice the strength to weight of aluminum,  but are even more vulnerable to overheating.  You cannot replace all the metal with all-composites,  except in minimum-velocity suborbital flight,  and even that is on a heat-sink transient. 

I hope these comments help provide additional guidance for those wishing to use my results.  I really do appreciate the comments,  Google +1's,  and other feedback.  Thanks,  and have some fun playing with this stuff.  I certainly did. 

--GW

More Gun Control? No Way!

I have a real problem with the disrespect for our constitution evident in the guest column written by Rosa Brooks (1),  that appeared Friday 9-20-13 in the Waco “Trib”.  On the other hand,  the guest column by Charles Krauthammer (2) that appeared the same day,  is something I very much agree with,  which is actually unusual for me,  as we have very different approaches to politics. 

Krauthammer’s article makes very good sense:  the mass shootings of late mostly seem to have been committed by deranged folks who “dropped through the cracks”,  either in the medical care system,  or in the gun background check process.  The scandal is not the violence they commit,  or that they use guns to commit it,  the scandal is that they drop through the cracks,  which tells us what really needs to be fixed. 

The Second Amendment

As for understanding the intent of the second amendment that Brooks so derides along with the entire constitution in her article,  I don’t need some court to tell me what it means. 

All I need is junior high grammar skills,  and a little junior high-level history for its context.  Here is the text,  as downloaded from Wikipedia (3).  It is a very short and straightforward item:

A well regulated militia being necessary to the security of a free state, the right of the people to keep and bear arms shall not be infringed.

There are only two phrases separated by a comma,  which version is what the states ratified when they ratified the constitution.  This was verified by its author,  Thomas Jefferson.  Other versions with extra commas and capitalizations are not what was ratified,  which means ascribing intent to the extra commas and capitalizations is bogus.  Period.  (Congress did that when they put it in the records.  Surprise,  surprise!)

The first phrase is a justification,  not a modifier.  Interpreting it any other way violates the basic rules of English grammar,  which have not really changed since the amendment was written.  The second phrase is a statement of what is to be done,  plain and simple.  The only “modifier” is not in the amendment,  it is in the context of the times (more about that below).

That first phrase calls forth the concept of a militia (as opposed to the “mob” that carried out the French revolution).  A militia was then,  and still is,  “citizen soldiers” that grab their weapons and come forth from home to fight on the battlefield in an emergency.  This was then,  and still is,  a different concept from the standing regular army (by extension today’s armed services).  

Historical Context

Back then,  the militia was the “Minute Men”.  Today’s real militia is the national guard,  and the inactive reserves that could be called back to duty.  In a really serious “nightmare-scenario” emergency,  the old “Minute Men” precedent still stands:  armed civilians could “spontaneously” join the ranks of the militia,  bringing their personal weapons with them.   Simple enough.

Also back then,  there was the concept of the armed population making the threat of armed revolution credible.  That was intended to make government behave itself,  instead of turning to dictatorship,  as so many others did before.  That risk,  and that need,  are still with us,  and in my opinion not enough “revolutionary threat” has been made in recent times:  government has been misbehaving!  (I think most might agree with that last assessment.) 

Also in historical context,  there was never any question that cannons and their ammunition would be kept in any other place than an arsenal.  That’s still true today,  and extends to missiles and a lot of other heavy-weapon stuff. 

Applying the Concepts

My point here is that both the regular army and the militia have to have the same kinds of weapons,  which back then were also exactly the same as civilian hunting weapons.  Our national guard uses pretty much the same weapons as the regular army,  although maybe not the “latest-and-greatest” versions. 

The only thing that has really changed is firearm technology:  there are repeating rifles,  machine guns,  and repeating handguns available today that were undreamed-of in Jefferson’s time.  The battlefield demands full-automatic weapons (machine guns) that in some cases (real assault rifles) can be operated as semi-automatic (one shot at a time without manual reload). 

If the “Minute Men” concept must ever be employed (and none of us wants that,  but prudence demands that we be prepared),  then the weapons brought from home need to be at least marginally useful on a battlefield with today’s machine guns.  That’s semi-automatic rifles and handguns that could (in an emergency) be quickly modified to operate as full-automatic weapons. 

There’s no way around that technical requirement.  You cannot ban semi-automatic weapons like the AR-15 (which is not a real assault rifle,  because it cannot operate as a machine gun),  and still have civilian citizen-soldiers credible for that last-ditch,  maximum-emergency battlefield scenario. 

What you do instead is ban that full-auto modification,  and then de-facto suspend it for that nightmare scenario.  We already did that,  with the 1934 firearms law that outlawed civilian possession of fully-automatic weapons.  It made sense then,  with the organized crime violence,  and still does today. 

What About Magazine Size?

As for the politically-popular idea of limiting magazine size,  it is easy to show by elementary calculation that switching-out large versus small magazines does not really affect firing rates very much,  averaged over multiple magazines.  So,  there’s not much to gained in mass-shooting scenarios by doing this. 

But,  if the citizen soldier’s magazines are interchangeable with the military’s,  then he is more effective in that last-ditch nightmare scenario because of vastly-improved logistics.  So,  limiting magazine size is a bad idea:  it doesn’t solve the civil violence problem,  but it does degrade our maximum militia capability.  Simple,  clear,  and compelling.

So,  What Do We Do?

If you want to reduce gun violence in America,  then do something about all the untreated mental illness all around us.  First on the list.

The main “leak” in the gun background check process is:  nothing may be done to prevent sales to the mentally ill until they have been judged mentally ill by a court. 

We have had a lot of mentally-ill mass shooters recently,  and none of them were so-judged by any court.  Fix THAT.  Make it a staged process,  instead of an either-or choice that only lawyers and judges control. 

Do those two things,  and you won’t need any of the other “gun control” ideas. 

Then you won’t have to degrade that maximum-defense capability that the civilian volunteer provides for that nightmare scenario,  or the citizen revolution capability that makes government behave itself.    

It ain’t rocket science,  as Brooks said in her article (the one thing I did agree with).  Nor do any of you have to be one,  in order to understand this issue,  or what to do about it. 

References:

(1) Rosa Brooks, "Over and Over and Over Again",  opinion-page article in Waco Tribune-Herald newspaper,  appearing Friday 9-20-13.

(2) Charles Krauthammer,  "Our Society Abandons Mentally Ill,  At Huge Price",  opinion-page article in Waco Tribune-Herald newspaper,  appearing Friday 9-20-13.

(3) results of google search for "second amendment,  being the article on Wikipedia,  as it appeared Friday,  9-20-13.  (This wording of the amendment matches that in my history books.)

Related Articles

I have written about this topic before,  with pretty much the same message,  just different details being addressed.  The other articles are listed here by date and title,  all on this site.  I added the search keyword "guns" to all 6 articles. 

2-5-13:  Real Problems with the Proposed Gun Control Legislation Items
12-20-12:  On the Tragedy in Connecticut
12-14-12:  School Shooting in Connecticut
8-9-12:  Mass Murder Shooters and Gun Control
1-13-11:  On the Shooting Rampage in Tucson

Update 9-22-13:

Since posting this article (or any of the older ones),  I have seen nothing in the news (or anywhere else) to induce me to revise these opinions.  Not even Wayne LaPierre's recent inflammatory remarks alter my views.  He was actually fundamentally correct,  in that there surely seemed to be a dearth of armed guards defending the otherwise gun-free zone that was a navy base facility. 

Update 12-15-13:

I have seen nothing to change any of my conclusions or recommendations in this,  or any of the referenced articles,  not even with the latest shooting incident at Arapahoe High School in Colorado.  In point of fact,  the Arapahoe HS incident confirms what I have been saying. 

The news reports indicate that the Arapahoe HS shooter killed himself only 80 seconds into the incident,  because he knew the on-site deputy was “on the way”.   This quite apparently stopped him from shooting more kids and/or teachers. 

My contention is that we should be defending our gun-free zones on a 60-second time scale,  just as what worked very well in the old west.  At Arapahoe HS,  unlike Newtown elementary in CT,  there was a deputy on site.  That so very obviously made a huge difference to the outcome. 

What is unclear about Arapahoe HS in CO is whether getting down from 80 sec to 60 sec might have saved the 17 year old Ms. Davis from her shotgun head wound.  But,  what is clear,  is that more did not die,  as was the shooter’s intent. 

Properly-defending declared gun-free zones really does work,  just like it did a century (and more) ago. 

We already know this works.  So,  just get on with it!

Update 5-28-14:

The latest incident with Elliot Rodger in California is simply another crazy never institutionalized,  and so still able to legally buy guns without any impediments at all.  That gets right back to the text highlighted in yellow in the main article above,  under "So What Do We Do?"

This is "lawyer/judge nonsense" at its very worst!  Until that loophole is closed with a policy that makes good common sense,  these horrific incidents will continue,  mark my words!  Nothing else is going to fix that very fundamental lack.

Reusable Chemical Mars Landing Boats Are Feasible

I was initially quite skeptical that a single-stage chemical lander could be feasible for one-shot use, much less reuse-for-multiple trips,  with refueling.  But,  after some investigation with bounding analyses,  I have changed my mind.  These things are very feasible.  So,  we need not incur the difficulties of solid core nuclear propulsion,  in order to have a very practical ferry capability between surface and orbit at Mars.

Propellant Options

I do not know what propellants might really turn out to be manufacturable on Mars.  Right now,  I doubt anyone else really does,  either.  So,  I picked four combinations that might be “typical” of things others are considering.  The exact species were simply what was available in some old references on my bookshelf.  The four,  with rationale,  are:

Liquid oxygen – rocket-grade kerosene,  LOX-RP1,  well-developed technology

Nitrogen tetroxide – unsymmetrical dimethyl hydrazine,  NTO-UDMH,  well-developed, fully storable

Liquid oxygen – liquid methane,  LOX-CH4,  new,  easily-manufacturable?

Liquid oxygen – liquid hydrogen,  LOX-LH2,  highest-performing,  from ice?

Rocket Ballistics and Design

The data I had on-hand in my library allowed me to determine an estimate of rocket characteristic velocity (c*) as a power function of chamber pressure for all four selections.  I was able to scale the c* to an easily-achieved design goal of 500 psia (3.447 Mbar) for a medium-to-small rocket engine.  Expanding from that,  to 6 mbar backpressure,  gave me expansion ratio and ideal thrust coefficient,  read off a standard chart.  I corrected that by an appropriate nozzle efficiency reflecting realistic half-angles,  for single-engine uninstalled engine performance and gross geometry, derivable from a thrust specification. 

Once sized by a thrust requirement,  the throat diameter is known,  which with the expansion ratio and half angle sizes-out the bell dimensions.  I picked an arbitrary 10:1 area contraction ratio chamber-to-throat,  and sized chamber length off some empirical L* values.  Thus “reasonable” engine dimensions can be estimated,  once a thrust requirement is set.  The point is “getting into the ballpark”.

The Payload

I did my study around the concept of a fixed payload to be landed.  For an early mission,  this might typically be 3 men with a month’s supplies.  I included a guessed allowance for a rover car with a drill rig on it,  and allowances for surface exploration and experimentation gear,  plus a small inflatable “pup tent” in case they were too far from the lander to return.  I got 3.191 metric tons to cover this,  but that’s just a guess.  So also is 60+ cubic meters to contain it all,  plus some “living space” inside the lander. 

Flight Paths

These vehicles will have ballistic coefficients far larger than any of the lander vehicles so far delivered to Mars.  Investigations I have run came to the same conclusions as others:  such vehicles can aerobrake successfully,  but will come out of hypersonics at too low an altitude for chutes or ballutes to deploy,  much less do any good,  in the thin “air” of Mars.  Therefore,  hypersonic/supersonic retro-propulsion is going to be mandatory.  The vehicle simple rocket-brakes directly to touchdown,  once the entry hypersonics are over,  or at least mostly-over.  That kind of descent is illustrated in Figure 1 (below). 

These estimates are based upon a surface-grazing transfer ellipse from a 200 km altitude orbit (low Mars orbit,  LMO).  The de-orbit burn requires only a 50 m/s “delta-vee”,  something that attitude thrusters can provide.  LMO speed is 3.455 km/s at 200 km.  Speed at the 140 km interface is 3.646 km/s.  Speed at the local Mach 3 point varies some,  but falls in the 0.7 km/s range.  Allowing for heavy rocket-braking inefficiencies plus final near-hover for touchdown,  a good guess for the effective ideal “delta-vee” for descent is 0.9 km/s,  after getting some 2.9 km/s deceleration from aerobraking. 

The ascent is easier to rough-estimate.  The total velocity to be achieved is LMO speed,  with empirical “kitties” added to cover gravity losses and drag losses.  For rough-estimating clean rockets here on Earth,  5% of target velocity is adequate for each of both kitties.  The losses are less at Mars because of the weaker gravity and thinner “air”.  I simply ratioed-down the 5% figures by the surface ratios of gravity and density,  to a combined-loss “kitty” of 1.94%.  The ascent ideal “delta-vee” figure is then at most about 3.6 km/s,  as given in Figure 2 below. 

Those “delta-vees” add (for no reduction in payload upon ascent) to a total two-way “delta-vee” capability of about 4.5 km/sec.  For estimated installed specific impulse performances,  excepting hydrogen,  the resulting propellant fractions then fall near 75%. 

The Vehicles

Each propellant combination leads to a different vehicle size-out.   My concepts are based around a guessed constant 20% inert weight,  to cover both structural and landing equipment items,  but hopefully with enough robustness to provide a significant service life in re-use.  Once a propellant fraction has been determined,  it and the inert are deducted from one,  determining the payload fraction.  The known payload weight then sets the entire weight statement.  The key is thus determining the propellant fraction,  from performance parameters and the total “delta-vee” requirement. 

My vehicle layout concept is based on historical US capsule shapes:  blunt heat shield with a more-or-less conical afterbody.  Heat shield shape is spherical-segment,  with a radius of curvature equal to the diameter.  I used a conical afterbody in the 20-to-30 degree half-angle range,  with a cylindrical extension,  since cylindrical tanks and pressure vessels are easier to build than conical ones.  The flight control station is at the end of the cylindrical section,  for really good visibility. 

The structure would be a deck frame with attached heat shield,  and four extendible landing legs.  Height-to-stance width is near one,  for good stability.  In the center is a cylindrical sealed compartment containing 4 canted engines (I chose 10 degrees arbitrarily) to “enforce” retro plume stability during descent.  With a sealed engine compartment,  there can be no throughflow through wide-open ports in the heat shield for the engines.  That eliminates the need for port covers,  and for swapping ends during hypersonic flight. 

A small makeup-massflow/coolant-flow might be needed to balance the volume-filling transient as the vehicle descends,  in order to prevent intrusion of any entry plasma through the open ports.  That was not analyzed here.  But,  it should be noted that this method of preventing throughflow (with a sealed engine compartment) should work equally well,  whether the engines are firing,  or not!

I arbitrarily picked an installed thrust sizing such that the vehicle accelerates at two standard gees,  at its maximum ignition weight.  I also assumed 4 engines in the cluster.  Correcting for cant angle and four engines,  produces the individual engine size-out,  and thus the necessary engine compartment dimensions,  and all the performance parameters. 

The required quantity of propellants and their densities sizes the tankage volume to stack on top of the engine compartment.  The 45-degree cone,  on top of all of that stack,  is the crew flight station.  The conical segment around the engine compartment is the crew living space plus cargo volume plus compartmentalization for airlock purposes:  60+ cubic meters.  I assumed that some of the conical shell panels are hinged at the deck line,  so that they could also be used as unload ramps.  Only the LOX-LH2 vehicle sized-out such that the conical segment had to extend partway up the cylindrical propellant tankage section,  a consequence of its lower mass (but similar overall volume) relative to the other three.  That vehicle had the smallest engine compartment,  by far. 

Figure 3 below shows rough vehicle layout and dimensions,  all four being roughly the same overall shapes to within a fraction of a meter.  The weight statements are quite different,  as a function of propellant selection,  as shown in Figure 4.  Note than none of these vehicles could ride assembled to Earth orbit inside existing payload shrouds,  which are around 5 m diameter max. 

These things will have to be assembled on-orbit in low Earth orbit (LEO) from docked and assembled components,  and then sent to Mars.  I would suggest each lander push its at-Mars propellant supply to Mars as an unmanned cluster vehicle,  via min-energy trajectory.  The quantity of propellants each lander pushes to Mars then depends upon the scope of planned on-orbit-based operations.  Determining that is out of scope here.

Rough Performance Estimates

These vehicles could be operated as orbit ferries in either of two ways:  (1) on-orbit basing,  meaning refueling from supplies in LMO,  and (2) surface-basing,  meaning refueling from supplies manufactured on the surface.  In the first case,  entry is at very nearly maximum mass,  followed by ascent at reduced mass.  In the second case,  ascent at maximum mass is followed by entry at much-reduced mass.  These two cases very much affect the entry ballistics,  resulting in much-different final rocket-braking needs. 

On-orbit basing is the more stringent descent case.  My analyses indicated rocket-braking requirements in the 3-to-5 gee range,  if braking was delayed to the local Mach 3 point.  This violated considerably the two-gee engine sizing assumption.  Indicated “hover time” at 200 m was barely adequate at 55 sec to cover the touchdown.  This outcome simply indicates that rocket braking must start earlier,  closer to the entry max deceleration gees point,  which is at substantially-hypersonic speeds.  There is no reason this could not be done,  once the hypersonic/supersonic retro-propulsion approach is adopted at all.  I did not re-analyze this change in detail,  having already established the basic feasibility. 

With surface-basing,  the descent rocket braking requirements (waiting to local Mach 3) all fell within the two-gee intended design.  The 200 m hover times were at least twice those of the other case,  for all four propellant options. 

For both cases and all four propellant options,  entry peak deceleration gees fell in the 0.71-0.73 range.  It is the 2-gee rocket braking that is the highest gees the crew must withstand during descent.  Accelerations upon ascent were not analyzed,  but should be comparable to 2 gees,  assuming throttleable engines.  Clearly,  the crew should arrive at Mars fully physically-fit,  and able to endure a few minutes at 2 gees,  seated,  with full human functionality. 

This implies that their transit vehicle should provide artificial gravity (by spin) at 1 full (Earth-normal) gee.  That is an issue for manned transit vehicle design,  out of scope here,  but very definitely needing attention called to it.   

Employment of the Landing Boats

These craft could be employed flying multiple missions,  either on-orbit based,  or surface-based.  It is easy to imagine a first manned exploration mission to Mars,  where the landing boats are operated on-orbit,  using propellant supplies brought from Earth.  At the end of the mission,  the landers and any remaining propellants would be left in Mars orbit “for subsequent missions to use”. 

A subsequent mission with the objective of establishing a more-or-less permanent base or outpost,  might use them differently.  The initial landings,  to ferry vehicles and equipment down,  might use on-orbit propellant,  with subsequent flights using propellants manufactured on the surface.  Once that capability exists,  and we are interested in only one (or maybe two) specific outpost sites,  the landers would be operated in surface-based mode,  probably for the remainder of their service lives. 

A variation on that scenario would be to accomplish both objectives in one manned mission to Mars,  something feasible because the stay time at Mars until the orbits are “right” for the return is over a year long.  Initial on-orbit based explorations are done,  until a site is identified where large quantities of propellants can be manufactured fairly rapidly.  Then the vehicles and equipment are all transferred to that site,  for surface-based operation,  from what will become a more-or-less permanent outpost.  This presupposes that propellant can be manufactured in 20-ton lots on a timescale of a few weeks,  on that first manned trip.  If that is not true,  we are inevitably reduced to the first scenario. 

I would suggest that we plan for the two-objectives-in-one-mission scenario,  but with sufficient on-orbit propellants to fully support the fallback position. 

About the Heat Shield

The entry analyses yielded a worst-case peak stagnation heating rate of roughly 5.5 W/sq.cm.  This is low enough to allow the use of black-surfaced low-density ceramic heat shield materials on the windward surfaces,  even at the stagnation point.  White-surfaced low-density ceramics can be used on all lateral and leeside surfaces. 

However,  because these vehicles land on natural regolith surfaces,  there is the dead certainty of suffering dirt and stone impacts to the heat shield,  due to rocket back-blast effects at takeoff and touchdown.  Low density ceramics far less fragile (and far less labor-intensive to maintain) than the well-known Shuttle tile are thus demanded. 

There might be one:  my oddball experimental ceramic-ceramic composite appears to have the necessary toughness,  with no “show-stoppers” anticipated to completing its development.  This material is described in the posting dated 3-18-13 and titled “Low-Density Non-Ablative Ceramic Heat Shields”.  My recent well-received paper at the 16^th Annual International Mars Society convention (in Boulder,  CO) also covered this same proposed material.   

Making the Propellant Selection

That topic is beyond scope here.  I do not know what propellant combinations might actually prove practical to manufacture on Mars.  I doubt anyone else really knows yet,  although many might claim to know.  That answer needs to be found first,  so that the landers we design and send,  on that first manned mission,  are compatible with what we can actually produce there.  That way,  they can serve for quite a while.  Making that selection is the fundamental pacing item for picking a lander design approach,  and then making it a flight-ready vehicle. 

Conclusions

1.       The first priority is to decide “for sure” what propellant selection could “best” be manufactured on Mars.

2.       Next,  the “landing boat” design given herein,  matching that propellant selection,  should serve as the design start-point for an actual “landing boat” design.  This design should be built,  tested,  developed,  and readied for operational use. 

3.       The manned Mars mission,  or sequence of missions,  should employ this “landing boat” in the relevant role (or roles),  and these craft should be left there for future uses. 

4.       This “landing boat” should be designed with the maximum possible expected service lifetime,  far beyond the needs of one mission,  so that it may serve subsequent missions or roles with local refueling,  for as long as possible. 

5.       There is no reason this “landing boat” development could not be started right now,  and thoroughly tested in low Earth orbit,  just as the lunar lander was.  The pacing item is (again) deciding which propellant selection could actually be manufactured upon Mars. 

 

Figure 1 – Descent Trajectory Assumptions and Requirements

Figure 2 – Ascent Trajectory Assumptions and Requirements

Figure 3 – Rough Overall Vehicle Dimensions,  Without Internal Layout Details

Figure 4 – Weight Statements and Performance Parameters by Propellant Selection







Update 9-9-13:  Some Further Thoughts About Selecting the Right Propellants

I really don't think kerosene is something we could practically manufacture on Mars,  but it might be somewhat representative of a hydrocarbon heavier than methane,  that we might dream up a process for.  It is a very well-known technology.  But,  I'd bet we can find a way to ignite or keep-unfrozen any of these choices,  though. 

I really don't think NTO or any of the hydrazines might actually be practically manufactured on Mars,  without a source of fixed nitrogen.  That's a huge obstacle there,  as far as I know.  But,  we already know those propellants can be easily stored,  and we have had engines that re-light multiple times in vacuum with them,  for decades now.  That's pretty much the technology of the shuttle OMS maneuver engine pods. 

I suspect LOX-LH2 would actually be the "easiest" to manufacture on Mars,  using mined ice and electrolysis as the basis.  LOX is not too much trouble to liquefy and store;  LH2 is much trickier to do,  with the ortho vs para form problem perhaps now the easiest problem of several to resolve. 

On Mars,  the truly fundamental problem is "where is the ice deposit big enough to be worth mining?"  We now know for sure that Mars has lots of water still (in the scientific sense),  the trouble is that it's just not "everywhere".  The kind of ice lenses Phoenix found near the pole is not the kind of deposit that supports practical mining and manufacture.  What we need is a buried glacier 10+ meters thick and many,  many km in lateral extent. 

BTW,  it'll be subliming as we dig it out.  Every mine hole has to be regolith-buried when not in use.  There will be one whale of a lot of regolith-moving operations involved in this activity.  The machines will look like heavy mining and road-building equipment.  That takes a big lander,  even if shipped in small pieces and assembled on site.  These things will not be carried by a series of Apollo-like dinky-little landing modules.  No way.  We need real "landing boats" of very significant size. 

They're not gonna fit existing payload shrouds for launch to LEO for this mission.  Something else to think about. 

We have orbital observations of where some such buried glaciers might be (emphasis on "might"),  but we have absolutely no ground truth about it.  I have never seen a robot probe design capable of determining that kind of ground truth,  either.  So,  if we are going to plan on making LOX-LH2 to return,  where do we land? 

Tough question.  We have to be close enough to walk to the ice,  or it ain't gonna work.  We're talking front end loaders,  bulldozers,  and large pressure-vessel process machinery here,  with maybe even some pick-and-shovel work by more than 2 men.  Long range transport is simply out of the question,  that first time up with propellant manufacture. 

As for making methane out of water,  and the CO2 in the "air",  the low inlet densities make all your machinery (whatever it is) look very large and heavy and energy-intensive,  compared to what we are used to here at home,  by about a factor of 14.  My guess is you can make 1's,  not 100's,  of kg per day.  You'll not accumulate enough to return a crew (tens of tons),  not even in a year's stay,  even if it doesn't break down or encounter unexpected problems. 

And you will encounter unexpected problems (lots of emphasis on "will").  Done robotically before the men arrive offers a potential way out,  except that robots-as-we-know-them-today are simply inept at solving unexpected problems.  Put the men there to solve those problems,  and you are right back to the inability to accumulate tons of propellants in time.  Plus,  with LOX-CH4 you still have to solve basically the same water problem as LOX-LH2,  to get the oxygen and the hydrogen. 

So I dunno which one to try.  And I don't yet see much of a path to resolving this in time for a mission in the 2030's,  much less the 2020's we'd all like to see.  NASA has no plans to send the right kind of probes that could locate the propellant-making resources.  I don't see anybody else sending the right kind of probes,  either.

That puts me back to the costly-but-sure-thing concept:  first mission relies on propellants-sent-from-Earth.  Which means it is an LMO-based mission,  sending down multiple ferries to multiple interesting sites,  and emplacing the machinery to experiment with propellant manufacture at the most promising ones after the men return home.  Leave the ferries in a higher Mars orbit,  with whatever propellant is left over,  for the next mission to use.  What's the point of going all that way with men,  and only making one landing?  That's really dumb!

Meanwhile,  we have to guess which propellants might actually be made on Mars most practically,  and build the first-mission ferries to use that.  That way subsequent missions (including planted bases) can refuel and re-use the same ferries.  Right now,  I'd guess LOX-LH2 from ice.  But with an engine compartment big enough to accommodate being refitted with different engines.  And with compartmentalized tankage to accommodate being re-plumbed for different propellants.  That's heavier,  and so is structural robustness necessary for long-term reusability.  My assumptions of inert structural weight 20% are quite likely too low. 

 Update 9-11-13:

I have had some conversations with John Strickland about his designs versus mine.  Coming from very different starting points,  his results for lander vehicles and mine are amazingly similar.  This study of mine is a lot more realistic than should ever be expected of a bounding calculation.  (And that's what it is,  so don't read too much into all the nitty gritty little details).   

Update 9-22-13:

I'm beginning to think,  since all 4 of my vehicle rough-outs turned out to be about the same overall physical size,  that a "good" design might be one based around LOX-LH2,  but with enough space internally to be reconfigured and re-engined for LOX-CH4.  Comments?  Ideas?  Please weigh-in!

My best guess is that this LOX-CH4 combination is the "most likely" in-situ-produced propellant combination,  long term.  Short term,  I really think it might be LOX-LH2 from mined ice,  with nothing but electrolysis and liquefaction.  That is the simplest and most direct combination we have.  It is only restricted by where significant ice is actually buried on Mars. 

Basically,  vehicle size is fundamentally "set" by the payload mass to be landed.  My concepts are for "well-empowered" explorers,  not permanent base builders.  Please weigh-in with payload mass ideas for the follow-on base-building missions.  I know a lot less about that. 

The real question to answer here is:  what are we really going to do with men on Mars?  Explore?  Build bases?  Both sequentially?  Both at once?  The answer makes a huge difference to the mission approach,  architecture,  and component designs.  Overwhelming,  actually.