Israel vs Hamas: It Is Worse Than You Think

It does not matter whether you are a Palestinian supporter or an Israeli supporter,  you are likely wrong in your beliefs,  precisely because you do not understand how complicated all of this really is.  There’s more than enough blame to go around for both sides,  and yet more blame besides that.

The figure below shows a map of the Palestine region as it was proposed to be divided between Jews and Palestinians by the UN’s 1947 partition proposal.  That was a 56-44 split of area between Jews and Palestinian Arabs,  with Jerusalem intended to be an open city belonging to neither side. While not an even split,  the Jewish area included a lot of the more-or-less uninhabitable Negev desert in the south.

The neighboring Arab nations did not agree to this,  but it was pretty well done anyway by the UN and western nations in 1948,  with the creation of Israel,  pretty much along the proposed lines pictured. 

Almost immediately,  the neighboring Arab nations invaded,  entering through the Palestinian territories,  starting the 1948 war,  which they ultimately lost,  losing some of the Palestinian territory to the Israelis,  as indicated in the figure,  but keeping the bulk of it as shown.  Although they could have,  these neighboring Arab states never gave the Palestinian territories they occupied back to the Palestinians!

Multiple Wars Since

There were wars in 1951-1955 (the Palestinian Fedayeen insurgency) and in 1956 (with Egypt over the Suez Canal) that did not materially affect the map shown in the figure.  It was essentially unchanged until the 1967 “6-Day War”,  when Israel captured the Golan heights in the north from Syria,  the Gaza Strip and Sinai Peninsula from Egypt,  and much of the West Bank from Jordan in the east.  The Sinai Peninsula was eventually returned to Egypt as part of the peace accords after the 1973 “Yom Kippur” war.     

1967-1970 saw the so-called “War of Attrition”,  and 1971-1983 saw the Palestinian insurgency in south Lebanon.  The map didn’t really change,  with the Gaza Strip and much of the West Bank occupied.  1973 saw the “Yom Kippur” war,  resulting in peace accords and recognition of Israel by Egypt. 

Since then,  there was the 1978 First South Lebanon conflict,  resulting in the Palestine Liberation Organization (PLO) being expelled from southern Lebanon.  (That’s the same PLO that is today the government of the West Bank.)

There was the 1982 First Lebanon War,  resulting in the PLO being expelled from all of Lebanon.  There was also the 1985-2000 South Lebanon conflict,  which resulted in Hezbollah being based permanently in Lebanon (and Syria). 

Then there was the 1987-1993 First Palestinian Intifada,  which the Israelis suppressed.  During it were the 1991 Iraqi rocket attacks on Israel,  which was a strategic failure for Iraq,  since it did not provoke a response.  2000-2004 was the Al Aqsa (or second) Intifada,  also suppressed.

Then in 2006 was the Second Lebanon War,  where the Israelis took on both Hezbollah and the Lebanese military fighting together.  That was inconclusive,  but with large combat losses.

All of that is part of the background to the current 2023 Israeli-Hamas war that ongoing,  with war also starting in the north with Hezbollah in Lebanon.

Related Stuff

Another part is the history of Israeli settlements in the West Bank.  While the territory was occupied,  it was never officially annexed,  because if it were,  the Palestinian Arabs there would add to those living within Israel as Israeli citizens,  and outnumber the Jewish citizens.  Not annexed,  those settlements may well be technically illegal,  depending upon your interpretations of international law and UN resolutions.

Yet another part of this is the far-right nature of the coalition that is Netanyahu’s government in Israel.  It was that way when he was Prime  Minister before,  and it is that way now that he is again.  When he was Prime Minister before,  that’s when the pace of planting settlements in the West Bank vastly accelerated.  Those settlers would by-and-large be of a similar far-right mindset,  just to choose to settle there,  the whole point being to effectively make that Israeli land,  no matter what.

He is Prime Minister again,  and so effectively in command of the Israeli military.  The “court reform” is him trying to pave the way to staying in ever-more-power,  permanently.  The settlers in the West Bank that he essentially put there,  are now going around killing their Palestinian sheep-herding neighbors in this latest war.  And the Israeli military in Gaza is using strategy and tactics that reflect an utter disregard for Palestinian civilian lives.  Armies often tend to reflect their commanders-in-chief. 

Why would anyone be surprised at that behavior,  given who and what Netanyahu is?

And,  no one should be surprised that the Arab neighbor nations are refusing to take in Palestinian refugees from the Gaza Strip,  given the fact they never gave back the lands in 1948.  That border crossing with Egypt is closed.  It has been difficult to get them to allow UN aid trucks in. 

What is surprising is that no one is reporting on all this ugly history and background! 

And don’t forget what Hamas did with its October 7^th attack of unprecedented scope and evil.  They took hostages,  killed with great barbarity including women and children,  and raped a lot of women.  Those are internationally-recognized war crimes,  as is hiding behind human shields. 

You should also be unsurprised at the civilian casualties that inevitably happen when you strike back at an enemy like that.  You have to go through that human shield to strike them.  It is inherent. And,  they tried to prevent their own civilians from evacuating,  preferring instead that they die serving as human shields. 

And THAT evil is the government of Gaza?  They do not care whether the people they govern live or die!

My Predictions

Hamas will fight until it is destroyed,  but at high losses to the Israeli military,  and utterly enormous losses of civilian Gazan lives.  Hezbollah is lots bigger and better equipped than Hamas,  and will continue the war with Israel in the north,  after Israel’s forces have been damaged by Hamas.

Update 1-12-2024:  these two predictions have already come true.  Which makes the following prediction even more crucial:

Both Hamas and Hezbollah are well-known to be terrorist proxy armies funded,  supplied,  and commanded by the government of Iran,  which has so far suffered precisely zero consequences for what its proxies have been doing.  This violence is not going to stop until they do suffer for what they have done. 

List of Iranian Terrorist Proxy Armies

Iran is a terrorist dictatorship run by fake mullahs misusing religion to “justify” what they do.  They run a sham democracy in which they can over-rule anything an elected body decides,  they suppress dissent violently,  and they are essentially propped-up in power by their private army the Revolutionary Guard.  Sounds an awful lot like Hitler and his SA stormtrooper private army in 1930’s Germany to me!

Iran also operates multiple terrorist proxy armies.

Per “Country Reports on Terrorism: Iran 2021” by Bureau of Counterterrorism at US Dept. of State,

retrieved from their site 19 November 2023 (highlighting is mine,  for this article):

I would also point out the recent attacks on commercial shipping in that same region.  Many of these are being conducted from Yemen by the Houthis,  and recently,  there was a direct attack from Iran itself by the Revolutionary Guard,  at a ship off the coast of India!

Wake Up and Do This Right

This violence will never end,  as long as Israel or anybody else just swats at these terrorist proxy armies.  It can only end when the terrorist government of Iran is overthrown.  I have nothing at all against the people of Iran,  they are good people.  But their terrorist government must go!  Or there can NEVER be peace in the Missile East!

You Will Not Like Hearing This,  Either:

Ukraine is the West’s proxy against the Russians.  That has dragged-on nearly 2 years.  It is time to do more than just send aid and money,  Russia has to lose,  and soon!  Which will likely cause regime change in Russia.  And that would be a good thing,  even if it did not affect things with China!

But this DOES affect China!  If Russia does not lose,  and soon,  Xi in China will be further emboldened by the West’s apparent weakness,  and will start World War 3 in the Pacific by invading Taiwan. 

And the longer this Ukraine thing drags on,  the more likely he will invade Taiwan anyway,  whether Russia wins or loses.

So just wake up and get on with it!

Proxy army war can get complicated,  can it not?

 

Christmas Funnies 2023

This is just for fun.  My wife found a funny on the internet that I simply must repost here.  The lights as hung in this city display are just too funny,  given the appropriate modified Christmas song to go with them. 

As many readers know,  I put up a yard display each year that we call “White Trash Christmas”.  It involves some plastic yard flamingoes lit up from inside,  as Santa’s reindeer,  a lit-up plastic Santa,  and something I call the “Iron Christmas Tree”.  There are two views of this posted below,  taken just after sun-down,  Saturday 12-16-2023.  I put a trash bag of recycled cans in the wagon to represent Santa’s magic bag.

There are 10 flamingoes.  Eight are arranged in pairs,  marked and ordered as the eight names from the Christmas tales.  Those would be “Dasher”,  Dancer”,  Prancer”,  and ”Vixen”,  followed by “Comet”, “Cupid”,  “Donner”,  and “Blitzen”.  The lone one out front is labeled “Rudolph”.  The lone one behind is labeled “Bambi”.  Yeah,  I cheated.  But I had 10 yard flamingoes. 

The Iron Tree is something I made many years ago.  It has a PVC pipe center core,  and angled “legs” made of rebar.  The lights get hoisted up and then “may-poled” around,  to cover it. 

The view from more-or-less the side shows this thing set up near my farm shop.  The view from behind shows our house across the driveway from the farm shop. 

By the way,  each flamingo has pipe cleaner “antlers” on his head. 

Update 12-18-2023:  We also put up this year's version of the missile toad. 

And just to let you know that not everything around here is abnormal,  our tree is up,  too.

 

Conspiracies Have Short Lifetimes

As most readers of this blog already know,  I have long denounced the conspiracy theory that there is some evil “deep state” in control of our government and others around the world.  This derives largely from Qanon and similar or related sites on the internet,  all very popular with right-wing and far-right readers.  While some aspects and pieces of this movement date earlier,  this became a fearless leader cult around Donald Trump during 2015,  as he began his run for the Presidency in the 2016 election. 

Key to this is supposedly secret information about forces led by Trump to depose this “deep state”.  This is to take place as a “storm” during which “deep state” people will be rounded up,  imprisoned,  perhaps tried or perhaps not,  but definitely executed.  Invariably,  the persons identified as targets of this “storm” are Hollywood elites,  Democrats all over,  and any Republicans who are not Trump cultists (referred to as RINO’s).  Also invariably,  cultists refer especially to Democrats as being evil enemies of the state,  worthy only of being killed,  which has inspired considerable “lone wolf” violence of late.  

To accomplish this requires overthrowing our democracy and replacing it with a Trump-led dictatorship.  That’s the secret!  Which is exactly why so many Trump cultists have become publicly known to be “authoritarians” (meaning advocates of a dictatorship in America).  It is why Trump himself has recently begun saying “dictatorship on only day one” if re-elected.  The only part of that which is a lie,  is the word “only”.  And given his track record over the decades for egregious lying,  that is no surprise at all!

Secrets are hard to keep,  and with conspiracies to keep them secret,  there are invariably some numbers of people who know about them.  It has always been the pattern in the past,  that eventually somebody talks and spills the secret.   Thus there is some sort of finite “lifetime” that such information,  and the conspiracy to keep it secret,  both have.

I ran across a news item on the PBS NewsHour web site that describes the modeling of such conspiracy lifetimes.  This was “How Many People Does It Take to Keep a Conspiracy Alive?”,  which I retrieved from their web site 7 December 2023.  It was apparently published there 15 February 2016,  but I missed it back then.  The article reports the published academic work of one David Robert Grimes at Oxford University,  who found an equation modeling this effect as lifetime-to-revelation versus number-of-people-involved.  He reportedly published this equation in the Journal PLOS ONE.

I did not track down his published article.  The PBS NewsHour article gave enough numerical data for me to just reverse-engineer an approximation to what Grimes apparently found and published.  I simply keyed the quoted data into a spreadsheet,  plotted and re-plotted the data multiple ways,  and attempted some curve-fit equations.  One of those worked rather well,  and is probably fairly close to the Grimes result.  I even found one “bad data point” item quoted in the PBS news article.  My work is depicted by the plots in Figure 1.  It is the log-log plot options that produced fittable curves.

Figure 1 – Reverse-Engineering a Correlation Equation From Quoted Data (Excluding “Bad Data Point”)

The modeling equation that worked best,  as indicated by the higher Pearson’s r-squared parameter,  is in the upper right plot of the figure.  It models the data quite well.  It is:

               Y = 4.4378 [log₁₀(no. of ppl)]^-1.235

               Life, yrs = 10^Y

Now,  apply this to the Trump-cult (Qanon) conspiracy,  whose “secret” is that they want a Trump-led dictatorship in the US,  and also that they want to kill all their opposition!  This formed about 8 years ago in 2015 during the campaign leading to the 2016 election.  Solving iteratively the equation-in-reverse for a 8-year life,  then no more than about 4262 persons could know about the secret,  and still have it remain hidden the 8 years since then.  That’s clearly not the case,  as there are millions of Trump supporters who voted for him in both 2016 and 2020.  At least some of them are cult believers.

Alternatively,  in those elections,  there were 80-something millions of Trump voters.  I picked 82 million as an estimate,  just to be the right magnitude only.  Assuming all of them are part of the Trump-cult conspiracy,  the equation says the expected lifetime before the secret gets revealed would be only about 2.2 years.  The assumption I made is not correct,  but using it helps bound the problem between these two results:  2.2 vs 8 years after the 2015 campaign.  The “right” answer must lie between these points.

In actuality,  the revelation that Trump and his advisors were integrally associated with the Qanon theory came out during his Presidential administration January 2017-January 2021.  Trump’s then-adviser retired General Flynn publicly advocated for Trump to replace our democracy with a dictatorship led by Trump,  so that the “deep state” could be “overthrown” and eliminated!  This really means all their opposition being rounded up and killed,  exactly like what happened when Adolf Hitler’s Nazis took over Germany.  This revelation happened about 4 years ago,  exactly between the bounds I just defined!

Using the equation iteratively-in-reverse again,  on the 4 years since the 2015 campaign to sometime in 2019,  it estimates the number of really hard-core,  extremist Trump-cult believers to be about 109,700.  While not millions,  that’s large enough to be very alarming,  as the pool from which right-wing extremist violence can most likely be expected to come.  And we have already seen it.

And,  we already have several Representatives and a few Senators who so very clearly come from this cult.  So,  why is our Congressional dysfunction a surprise?  It’s part of their plan!  Chaos instead of governing helps to “justify” their radical change to a dictatorship under Trump.

Subtracting that roughly 110,000 true-believer cultists from about 82 million evident Trump voters,  that says something like 81.9 million voters have been deceived by this cult,  into voting for a would-be dictator over the US,  who has already attempted an insurrection and takeover on 6 January 2021. 

And THAT ought to scare the ever-loving shit out of almost anybody!

Please wake up out there!

Overall Study Results: Propellant From Moon

Just the overall results are given here.  The details supporting this are much more voluminous.  The basic notion,  Figure 1,  is to manufacture propellants on the moon,  probably using potentially-recoverable ice deposits near the south pole.  I initially looked at elongated halo orbits about the moon as a means to more easily access the south polar region off the moon from orbit.  Delivery is to low Earth orbit,  at low inclination eastward.  The idea is to base a lander on the moon,  flying loaded to a station in halo orbit,  and returning to the lunar surface with tank empties plus a bit of cargo.  The orbit-to-orbit transport vehicle to LEO,  is based at the station in the halo orbit about the moon.  It flies fully laden from there to LEO,  and returns to the halo station with empties plus that same little bit of cargo.  

Figure 1 – Basic Notion of Lunar Propellant Manufacture for Use in LEO

There are actually two halo orbit cases to consider,  although overall,  they are more-or-less a wash.  One is the proposed halo orbit for NASA’s “Gateway” station about the moon.  This one has an apoapsis radius that lies beyond the Hill sphere for orbit stability about the moon.  It also has a periapsis altitude much higher than that used during Apollo,  which acts to increase the delta-vee (dV) required of any lander operating from that halo station.  Long-term,  anything in this halo will eventually leave the moon and go into orbit about the Earth.  That is the direct result of having an apoapsis distance outside the Hill sphere.  Thus the halo station requires periodic correction burns,  just to stay in this orbit.

Using LOX-LCH4 propulsion that does not push the state-of-the-art as hard as SpaceX does with its Raptor engines,  it takes almost 48 metric tons of propellant manufacture on the lunar surface to deliver 1 metric ton of propellant to LEO via the NASA halo station,  and have the two vehicles return to their bases.  This does not include the propellant necessary for orbital correction burns to stabilize the station in this NASA halo orbit!    The transport vehicle is the smaller for this case,  while the lander is the larger,  as indicated in Figure 2.  This unstable halo orbit choice may actually be driven by the dV capability of SLS/Orion block 1 configuration,  which cannot reprise the Apollo 8 mission.

The “recommended halo” eliminates entirely the need to ship propellant to the halo station for periodic correction burns,  precisely because it is stable.  It has an apoapsis right at the Hill sphere limit,  and a periapsis at a distance comparable to the old Apollo missions,  to reduce lander dV requirements.  It takes about 51 metric tons of propellant manufacture on the moon to deliver 1 metric ton of propellant to LEO,  as also indicated in Figure 2.  For this case the transport is larger,  and the lander is smaller.  

Neither of these halo-based options is more “economical” than the projections for shipping propellant up to LEO from Earth’s surface,  for SpaceX’s Starship/Superheavy vehicle,  as indicated in the Figure.  That vehicle also uses LOX-LCH4 propellants,  so that comparison truly is “fair” in that sense.  

Figure 2 – Results for the Halo Orbit Cases

The alternative would be to manufacture propellant on the lunar surface,  and send it directly from there to LEO,  without utilizing any sort of halo orbit station as a waypoint.  This does require entry from the lunar transfer orbit directly into a low lunar circular orbit that is polar,  so that the south pole can be reached directly.  The reverse is the return.  Landings and takeoffs would be from-and-to this low polar lunar orbit. 

The dV requirements for the direct trip are higher than even the halo transport dV’s.  It was not possible to get “reasonable” results with LOX-LCH4 propulsion for this mission,  the Isp levels are simply too low.  I had to resort to the higher Isp of LOX-LH2 propulsion to get something “reasonable” in size.  However,  since there may well be recoverable ice,  but no free carbon,  on the moon,  it is far more likely that it will be LOX-LH2 propellants that actually get manufactured there,  anyway.  The corresponding results are given in Figure 3.

Figure 3 – Results For Direct Delivery of Lunar Propellant to LEO,  Using LOX-LH2 Propulsion

This is the only outcome that is better than trying to ship propellant up from Earth’s surface to LEO,  even using a vehicle as capable as SpaceX’s Starship/Superheavy.   It takes 14 metric tons of lunar propellant manufacture to support the delivery of 1 metric ton of propellant from the moon to LEO.  This includes the vehicle returning all the way to a landing near the moon’s south pole,  with empties and some cargo. 

Outcomes

In summary,  direct shipment from the moon to LEO is the best option,  but it will require the higher Isp of LOX-LH2 propulsion!  Using LOX-LCH4 is not feasible in any “reasonable” vehicle sizes,  primarily limited by achievable mass ratio as you increase propellant.  This requires manufacture on the moon of 25.1 metric tons of LOX-LH2 to deliver 1 ton of LOX-LH2 as payload to LEO. (corrected)

Failing that,  shipping propellant up from Earth’s surface to LEO is the next best option,  using the SpaceX Starship/Superheavy vehicle,  which is powered by LOX-LCH4 propulsion.  This is projected to require the manufacture of some 32 to 47 metric tons of LOX-LCH4 propellant on Earth to deliver 1 metric ton of it (or instead a ton of LOX-LH2) to LEO. This depends upon deliverable payload being 100-150 tons.

The halo station options turned out to be the least attractive.  Doing it with two vehicles instead of one reduces payload delivered for the propellant used,  by two vehicle payload fractions compounded (because there are two vehicles),  instead of just one.   The two halo options are not much different at about 48 (NASA) and about 51 (recommended) metric tons propellant to be manufactured on the moon to deliver 1 ton of propellant to LEO.  The NASA halo requires still more propellant manufacture on the moon than that about-48 figure,  because its halo station requires correction burns just to stay in its orbit long term!  The “recommended halo” avoids that requirement,  and gets a smaller lander design,  at the expense of a larger transport vehicle to LEO.

Conclusion

The final result says go with the higher-Isp LOX-LH2 propulsion,  and operate direct from the lunar surface to LEO fully laden,   and back to the lunar surface very lightly laden.  Return trips return the empty tanks for the next propellant shipment,  plus in this study,  a couple of tons of payload that could be base operating supplies.

Be Aware

The dV requirements used in this study include 8% of midpoint speed as dV budgets for course correction,  and 0.2 km/s budgets for rendezvous and docking,  once close (within at most 10 km).  Lunar takeoff ideal dV values were factored-up by 1.0083 for gravity losses,  and lunar landing ideal dV values were factored-up by 1.50 for losses plus the dominant hover and divert budget requirements.  Everything else was presumed loss-free “impulsive”,  for factor 1.00 applied to ideal values from basic 2-body orbital mechanics.  The 3-body approach and departure problems were approximated by the “far V” versus “near V” energy approximation. 

Explanations

There are some fundamental trends of mass ratio capability and the dV it can produce,  vs added propellant and Isp,  which help explain these results.  These are depicted generically in Figure 4.  

Figure 4 – Trends That Explain Results

The “knees” in these curves are not so apparent in the MR vs added Wp plot,  but they are very apparent in the dV vs added Wp plots parametric on Isp.  Where the slope is steep,  you will get a better (lower) propellant burned vs payload delivered ratio.  Where the slope is shallow,  that ratio will be large and unfavorable. 

Basically,  there is a favorable range of deliverable dV at each Isp level.  For the 450 s Isp level typical of LOX-LH2 propulsion,  this is up to 8 km/s dV for really nice results,  and up to about 11 km/s at the very most.  Beyond that,  it is a very serious diminishing-returns problem:  adding a lot of propellant for almost no improvement. 

For the 350-400 s Isp level typical of LOX-LCH4 propulsion,  the most favorable range is up to about 7 km/s dV.  At the most,  it is about 9 km/s.  Beyond that,  this is pretty much pointless.

By switching the halo station based lander and transport designs to LOX-LH2 (instead of LOX-LCH4),  the propellant to payload ratios could be significantly reduced,  perhaps looking more like those of Earth launch with Starship/Superheavy,  or maybe even slightly better. 

Regardless,  the lunar surface based and launched single direct transfer design approach is still the best,  despite it being only marginally favorable on the dV vs Wp curve. That is because it is a single vehicle,  and not two vehicles,  as in the other scenarios.  This scenario would look even better if its propulsion were nuclear thermal at 700+ s of Isp. 

As it is with LOX-LH2 propulsion,  the total vehicle dV requirement could be reduced a little,  making the propellant used/propellant delivered ratio even better,  if an LEO tug were used to retrieve the vehicle from an elliptical capture orbit about the Earth.  The same tug could put the vehicle back into the elliptical orbit for departure.  That reduces the arrival and departure dV’s significantly,  and it eliminates the rendezvous and docking dV requirement.  This gain is largely offset by the need for propellant deliveries to power the tug,  though.

Caveat

Ullage solutions for multiple burns with cryogenic liquid propulsion were NOT determined for any of these design rough-outs.  But they will have to be,  to flesh out all the design requirements!  Attitude control was also not addressed,  although given adequate acceleration levels,  some of those thrusters could supply the ullage function.  That is determined by the settling time constants that are acceptable. 

Corrections 12-10-2023:

I had not followed through fully on the spreadsheet for the direct vehicle.  The 14:1 delivery ratio figure goes with an otherwise-converged design that had far-insufficient thrust to takeoff and land,  resulting in too low an inert mass.    When I corrected that,  the vehicle proved to be enormous at 3000 tons,  with a really bad-looking delivery ratio.  I reconverged multiple times with multiple candidate engine numbers and thrusts,  until all the gees looked good,  including landing with hover capability,  but with takeoff reduced to 0.5 gees over lunar surface gravity.  That got me to the corrected figure.

How the Suborbital “Hopper” Calculations Were Made and with What

The Mars rocket hopper design rough-out was done using the course materials and tools for the “Orbit Basics +” course offered on the New Mars forums.  There is an “orbit basics” spreadsheet that does elliptical orbit 2-body calculations for either the two-endpoints case or the R-V-q observation case.  That tool’s R-V-q option can create suborbital trajectories,  which was done for the rocket hopper.  The spreadsheet calculates speeds V at periapsis,  apoapsis,  and at any one user-input radius. See Fig. 1. 

Figure 1 – The Two Cases Handled by the Orbit Basics Spreadsheet

To use this tool for the rocket hopper,  the most effective way was to define an exit (and by symmetry entry) point at the edge of Mars’s atmosphere,  and investigate various speeds V and exit angles relative to local horizontal.  Not every combination is allowable,  only certain values produce survivable peak heating and peak deceleration gees,  and also a feasible end-of-hypersonics altitude,  for a direct rocket-braked landing.  In fact,  many combinations produced instead a surface impact while still quite hypersonic,  in Mars’s thin atmosphere.

The symmetry of the exposed portion of the ellipse makes the V and angle “a” values the same for exit and entry,  at the entry interface altitude.  That is exactly how the suborbital trajectory analysis links directly to the hypersonic entry analysis. 

For the launch speed required of the hopper,  we need the speed along the orbit at the surface of the planet.  We need to be moving that fast at just about the same angle “a”,  at the end of the launch burn.  That is the theoretical dV^o value,  which needs to be factored up by about 1.02 to cover gravity and drag losses on Mars.  The factored-up launch dV is the mass ratio-effective value needed for proper use in the rocket equation.  See Fig. 2.

Figure 2 – Using the R-V-q Option in the Orbit Basics Spreadsheet for Suborbital Trajectories

The ”Orbits +” course covers launch,  entry,  descent-and-landing,  use of the rocket equation,  and estimating real engine performance,  as well as 2-body orbital mechanics of elliptical orbits. For the rocket hopper,  both entry and descent-and-landing apply,  using the methods and tools that are part of the course.  The direct rocket-braked landing is so simple,  it can be estimated from hand calculations.

The entry analysis is a 2-D Cartesian simplified analysis dating to about 1953,  and attributed to H. Julian Allen.  It was used in the 1950’s for estimating entry conditions for ICBM and IRBM warheads.  It was declassified by the mid-1960’s,  and then taught in engineering school classes.  Entry is presumed to happen along a straight line trajectory at a fixed entry angle.  The range is a crude estimate that you must wrap around the curved surface of the planet.  The constant angle you have to presume is relative to local horizontal,  as you move around the curve of the planet’s surface. 

These crude estimates get you “into the ballpark” only!  There is no substitute for a real digital trajectory program in polar coordinates,  but you do have to expend the significant efforts to construct the model to run in it.  At this stage of the game,  that is very inconvenient,  since the model to be input changes drastically as you iterate configurations.  Hence the need for a quicker ballpark estimate.

There is a lesson in the “Orbits +” course that deals with using the simplified entry analysis as a spreadsheet model of the entry process.  That spreadsheet is supplied as part of the course materials. 

In reality,  there is significant trajectory “droop” after the peak deceleration gees point,  that the simplified analysis does not account for.  I merely presume the local angle has increased to 45 degrees down,  by the Mach 3 end-of-hypersonics point,  when I do the rocket-braking by-hand calculations. 

There is also a lesson in the “Orbits +” course that deals with multiple ways to land after the hypersonics are over.  There is no spreadsheet,  but all the calculation equations are there to estimate any of these things by hand.  For the thin atmosphere of Mars,  from inevitable very low end-of-hypersonics altitudes with multi-ton vehicles,  there really is only direct rocket braking as a feasible thing to do. 

There is no time to deploy a chute,  much less get any deceleration from it,  plus there are no chute designs capable of surviving opening at Mach 3.  Even the ringsail chute designs used for probes at Mars have a maximum opening speed of Mach 2.5,  and slower-still is preferred as more reliable. 

Direct rocket braking is actually the simplest case,  and easily figured with nothing more than the simple kinematics of a high school-level physics course.  See Fig. 3.

Figure 3 – The Entry Model,  Plus Descent-and-Landing for Direct Rocket Braking

The vehicle layout and dimensions,  plus its weight statement,  are essentially custom hand calculations,  the suite of which is different for each different configuration class.  I started with three configurations,  but only one gave me the low ballistic coefficient that the entry analyses said I must have.  I included wide-stance folding landing legs for rough-field operations.  Clearly,  there are a lot of considerations to address.  I created a custom spreadsheet to estimate all these quantities rapidly,  since I had to iterate multiple times before identifying a feasible solution.

The “Orbits +” course has a lesson on vehicle layout,  and a spreadsheet by which to set the weight statement,  but that spreadsheet was not really suitable for this very specialized suborbital vehicle,  especially since it must enter the atmosphere,  and also do that entry dead-broadside to get the necessary lower ballistic coefficient.  It is critical to select the correct diameter for this kind of vehicle,  so that the lengths are in the correct range,  and those results must be compatible and consistent with the seating arrangements in the passenger cabin.  That’s why I did it as a custom calculation,  and why I created my own spreadsheet for that purpose.  See Fig. 4.

Figure 4 – Downselecting to One Configuration for Vehicle Layout

All of this is aimed at using the rocket equation to relate vehicle weight statement to its velocity-increment (dV) performance capability.  The spreadsheet in the lesson on vehicle sizing of the “Orbits +” course does exactly that,  in a spreadsheet that is supplied as part of the course materials.  Since I did the hopper with a custom layout sheet,  I had to include this rocket equation stuff in it. 

The classic rocket equation dV = Vex LN(MR) uses the vehicle weight statement (from a vehicle layout process) to determine mass ratio MR = Wign/Wbo,  and an estimate of engine Isp to determine the effective exhaust velocity Vex = Isp * gc.  It then gives you the performance estimate dV,  which must cover the mission needs plus any gravity and drag losses,  or other considerations,  such as hover and divert during landings.

There is a restriction on this:  you may sum the dV values estimated for all the mission burns into an overall mission dV,  only if the weight statement does not change between burns.  That means the payload and inert masses do not change,  and the only propellant mass changes are those for the burns. Failing that restriction,  you have a slightly different weight statement each time one of those items changes.  You must do a separate rocket equation calculation for only the burn associated with each slightly-different weight statement.  This hopper does not change its weight statement between burns!

For sizing vehicles,  the reverse process is what we really want to do,  for which the rocket equation rearranges to MR = exp(dV/Vex).  The engine Isp estimate gets us a Vex as before.  The mission dV is as before.  The layout gets us a payload mass and an estimate of vehicle inert mass fraction.  We use the rocket equation in reverse with the mission dV and the engine Vex to determine the MR that is required. 

This MR result determines the propellant mass fraction = 1 – 1/MR.  The payload fraction is 1 – propellant fraction – inert fraction.  Payload divided by payload fraction is the ignition mass,  ignition mass times the inert fraction is the inert mass,  and propellant fraction times ignition mass is the propellant mass.  Payload plus inert is burnout mass,  and burnout plus propellant is ignition mass.  In effect,  we are finding the vehicle weight statement from mission dV and engine performance to complete the vehicle layout process.  See Fig. 5.

Figure 5 – Using the Rocket Equation Properly to Size Vehicles to Missions

Clearly,  an accurate estimate of expected engine performance (as Isp or Vex) is crucial to the results!  There are a lot of references out there that list tables of Isp versus propellant combinations.  Just picking one right out of such tables is a serious error!  That is because engine Isp depends at least as much on the nozzle expansion characteristics,  as it does the propellant combination.  The expansion in the table is rarely the one you want to use,  and nozzle efficiency effects are never included in those tables. 

These things are all functions of the chamber pressure,  as measured at the nozzle entrance.  The chamber pressure value used in the tables is rarely the value you want to use. 

Finally,  Isp is directly affected by the engine cycle (through the dumped bleed gas fraction),  which those tables never include.  You can easily be 10%-or-more wrong just pulling values out of those tables.  Due to the exponential nature of the rocket equation,  that error in Isp can lead to fatal errors in your vehicle results for mass ratio and weight statement.

Thrust is often represented in terms of chamber pressure as F_th = C_F Pc At.  Isp is thrust divided by flow rate,  but it has to be the flow rate drawn from the tanks to be consistent with the rocket equation.  Flow rate from tanks = flow rate through nozzle + flow dumped overboard.  The flow rate through the nozzle relates to chamber pressure and c*-velocity as Pc C_D At g_c / c*.  And c* is a weak power function of Pc,  where the exponent is usually in the vicinity of 0.01.  The specific heat ratio of most rocket gases is in the vicinity of 1.20.  See Fig. 6,  for which the only propellant combination-related item is c*.

Figure 6 – How Engine Performance Must Really Be Estimated for a Specific Design

You are not totally free to set an arbitrary expansion ratio Ae/At!  It does not matter whether your nozzle is a “sea level” design or a ”vacuum-adapted” design,  any engine that is to be tested in the open air at sea level on Earth must not be allowed to flow-separate,  because that risks destruction of at least the nozzle exit bell!  Testing into a vacuum tank is extremely expensive!

For any given expanded pressure in the exit plane,  there is a value of the ambient atmospheric “back pressure” Pback that is “too much”,  causing flow separation.  That level is denoted Psep,  and it is easily estimated from the nozzle expansion pressure ratio:  Psep/Pc = (1.5 Pe/Pc)^0.8333,  which is an entirely empirical correlation developed for conical nozzles,  and is only slightly conservative for curved bells. 

For a “sea level” nozzle design,  you want predicted Psep = sea level barometric,  at some part-throttle Pc.  That way,  you can test in the open air for all power settings that high,  or higher.  The same is true of “vacuum-adapted” designs,  unless you give up testing in the open air!  But even then,  the engine and its nozzle still have to fit within the allotted space behind the stage.

The “Orbits +” course has a lesson on this topic,  plus a spreadsheet tool that does all these things.  It includes a database of c* and r-value data versus several propellant combinations,  as functions of Pc.

Updated 11-21-2023:  These very same methods were used to compute revised data for the upgraded Mars rocket “hopper” that could also serve as a personnel taxi to low Mars orbit. 

The original suborbital rocket “hopper” design summary was posted on this site as “Rocket Hopper for Mars Planetary Transportation”,  dated 1 November 2023.  The upgraded “hopper” that can also serve as an orbital taxi is posted on this site as “Upgraded Rocket Hopper as Orbital Taxi”,  dated 21 November 2023. 

There is a completely unrelated posting that deals with long-distance bulk freight transport on the surface of Mars.  That one is “Surface Freight Transport on Mars”,  dated 4 November 2023.

The final landing choice not described here is the lifting pull-up proposed by SpaceX for landing its Starship vehicle on Mars.  That is distinct from direct rocket braking,  and from parachute-assisted descents,  which require terminal rocket braking on Mars.  It is covered in the entry,  descent,  and landing lesson of the “orbits +” course materials. 

I did not examine that choice for any of these rocket “hopper” designs,  because I did not believe that my cylindrical layout has the mild-supersonic lift/drag ratio necessary to execute an aerodynamic pull-up,  even at very low altitudes on Mars.  I don’t really believe SpaceX’s Starship can do that either,  but that would be another study. 

To access the “orbits +” course materials,  which includes the spreadsheets,  go to the Mars Society’s New Mars forums online.  Go to the “Acheron Labs” section,  “interplanetary transportation” topic.  On about the second page of the list of conversation threads,  look for the “orbital mechanics class traditional” thread.  The course materials are actually posted elsewhere online,  but the links to each class session’s materials are in posts 3-to-21 of that thread. 

You will have to download the Excel spreadsheet files to make them functional.  The classes have a sort of lecture session (numbered) and a problem-working session (numbered with a “B” suffix).  These are available as Powerpoint slide sets and as pdf documents that are basically the traditional-style textbooks.  I recommend you download the pdf textbooks,  because all the explanations are in there.  They would be partly missing in the slide sets. 

Upgraded Rocket Hopper as Orbit Taxi

This article is about modifying a pre-existing design rough-out for a suborbital Mars rocket “hopper”,  into a design also capable of operating as a low Mars orbit personnel taxi.  That original rocket hopper design rough-out is covered in the article titled “Rocket Hopper for Mars Planetary Transportation”,  dated 1 November 2023,  and posted on this site.

 

               The Problem

Started with a suborbital “hopper”

               10 persons aboard on p-suits

               Short-term life support plus small luggage

Could it also serve as a low orbit taxi?

               Same payload

 

As indicated in the table just above,  I started with the earlier design rough-out that was only a suborbital “hopper”.  The idea was to carry 10 persons as the payload.  Although the cabin is pressurized,  these persons ride in pressure suits for a safety backup.  There are limited supplies of oxygen and drinking water,  plus minimal snack foods,  for up to a few hours’ ride.  A small luggage allowance was included.  The same payload would be carried to any low orbit destination.

As indicated in Figure 1 just below,  the suborbital trajectory is actually an ellipse in polar coordinates,  one with its periapsis inside the planet.  The vehicle launches into a gravity turn that reaches a suitable velocity and path angle at the entry interface altitude,  coasting from there. 

The best place to do a course correction is the apoapsis outside the sensible atmosphere,  where speeds are lowest and directions are easiest to change.  The entry conditions mirror the exit conditions,  with no burn.  The landing is a direct rocket-braked descent from the end-of-hypersonics point at local Mach 3 (about 0.7 km/s speed). 45 degrees of trajectory “droop” along a straight-line path is presumed.  I factored-up the speed to “kill” by 2,  to budget the final landing mass ratio-effective delta-vee (dV).

As illustrated in Figure 2 just below,  I used a surface-grazing ellipse as the initial transfer trajectory to the 300 km nominal low orbit altitude.  Like the long-range suborbital mission,  the vehicle launches into a gravity turn,  putting it onto the proper path at the entry interface altitude,  at end of launch burn.  Only the path angle is different,  being a lot smaller.  The entry point after de-orbiting is the mirror image. 

There is a small burn at apoapsis to raise the periapsis to the entry interface altitude,  with a period shorter than the target low circular orbit altitude.  This ellipse is the parking orbit in which to “chase” any target in the low circular orbit.  Once synchronized,  there is another small burn to circularize,  followed by a traverse to rendezvous,  plus a budget to actually dock.  Deorbiting is another small burn,  back onto the surface-grazing ellipse that guarantees entry.  The direct rocket-braked landing is identical to that of the long suborbital trajectory,  except that,  as it turned out,  the end-of-hypersonics altitude is higher,  coming back from orbit at the lower entry angle. 

Figure 1 – Suborbital Missions,  Longest-Range Shown  

Figure 2 – The Orbital Mission,  Including “Chase”,  Rendezvous,  and Docking  

To accommodate the more demanding mission,  I resized the candidate LOX-LCH4 engine design,  and revised the inert masses upward a little.  Entry conditions forced me to increase the diameter and length a little,  in order to keep the entry ballistic coefficient down to tolerable values.  The original rough-out had two sets of tanks:  mains and headers.   The landing and course correction propellant was in the headers,  with the launch propellant in the mains.  

This became 3 sets of tanks and two different engine designs.  The launch-and-landing main engines stayed about the same at 30,000 lb thrust,  each of 4,  drawing from the mains for launch and headers for landing.  I was able to increase the expansion ratio and specific impulse a little bit.  See Figure 3 for the basic layout revisions.

But course correction suborbitally,  and all the orbital maneuvering,  rendezvous,  and docking,  really needed much lower thrust levels.  So I sized some lower-pressure,  pressure-fed engines of only 550 lb thrust,  each of 4.  These used a small third set of 800 psi pressurized propellant tanks,  plus a supply of dry nitrogen gas at 2200 psi to power this,  in one of two options examined.  

Figure 3 – Revised Internal Layout at Larger Diameter and Length  

Because the inert fraction increased a bit,  I resized the expansion of the main engines to increase specific impulse a bit,  to compensate as much as possible.  The original “hopper” main engines had an expansion ratio sized for incipient separation at 67% chamber pressure,  if fired in the open air at sea level on Earth.  I raised that to 80%.  See Figure 4 just below.

The idea was to enable easy and relatively inexpensive development testing on Earth.  The change was small,  but every little bit helps.  These being full flow cycle,  turbo-pumped engines of significant thrust level,  I did not want to complicate things by adding vacuum bell extensions that were not regeneratively cooled.  These do not push the state of the art very hard,  being only 2000 psia chamber pressure.

Figure 4 – Reworked Main Engine Design for Slightly Higher Expansion 

The original “hopper” design study convinced me that I did not need the large main engine thrust levels to do course corrections on the suborbital missions,  or orbital maneuvering,  rendezvous,  and docking,  on the orbital mission.  I kept the redundancy of 4 engines,  but sized for crudely only 0.1 gee of vehicle acceleration,  once exoatmospheric. 

Since the propellant quantities would be small,  the simplification of a pressure-fed design would be beneficial.  Alternatively,  since the engines were small,  they could be fed by electric-driven positive-displacement pumps.  Either way,  I picked a simple conical bell shape,  two-piece,  with a bell extension that was not regeneratively cooled,  as shown in Figure 5 just below.  Development testing on Earth could be done without the extension,  but operations on Mars or in space would use the bell extension.  This was not a throttleable design.

I ran numbers both ways for the propellant feed to the maneuvering engines.  I did not like the pumping power required for the positive-displacement pumped option.  It implied very heavy batteries,  even for the modest propellant quantities.  Regulated constant inert gas pressure on the propellant tanks turned out to be the better option.  These used a small third set of 800 psi pressurized propellant tanks,  plus a supply of dry nitrogen gas at 2200 psi to power this.  The chamber pressures were low enough to keep the pressures fairly modest on the tanks,  so that at small size,  they were not that big an inert weight penalty.  See Figure 6 just below. 

There were many false starts and iteration cycles to achieve all of this,  none of which is covered here.  The result is summarized in the unavoidably-busy figure,  Figure 7 just below,  which includes a weight statement that also displays mass fractions.  

Figure 5 – Smaller Maneuvering Engines as Sized   

Figure 6 – Of the Options,  Pressurized Tanks Seemed Best  

Figure 7 – Summary Data for the Final Rough-Out Design 

Of interest would be the various tank volumes.  Bear in mind these are fully filled for the mission to low circular Mars orbit at only low inclinations eastward,  and also fully-filled for the long-range suborbital mission (at low or high inclinations).  The headers and maneuvering tanks are always fully-filled,  but the mains are only partially filled for the shorter-range suborbital missions,  so that entry mass is not too big.

Suborbital ranges from 9400 to just under 500 km were examined in this study.   Their entry angles turned out to be a strong function of the suborbital mission ranges.  All of those suborbital entry angles were considerably steeper than the return from the orbital mission.  They were determined by feasible altitudes at end-of-hypersonics,  and by feasible peak entry gee values.

Figure 8 just below shows the final plots I got of various flight data during entry and final descent.  The Suborbital trajectories form trends,  and the orbital data fall way off those trends.  Upper left is end-of-hypersonics altitude and entry angle vs entry speed.  Upper right are the peak heating values during entry.  Lower left are the trends of peak entry gee,  and average gee during the final rocket-braked landings.  There is a numbered key relating the missions to each data point in each plot.  No gee level exceeds 4.5,  and no stagnation heating level exceeds 12.5 W/cm².  

Figure 8 – Data for Entry,  Descent,  and Landing (E, D, & L)  

Once again in Figure 9 just below,  the suborbital data for surface temperatures also form trends versus entry speed,  with the orbital data falling far off of those trend lines,  plus a numbered mission key.  There are trends for surface temperatures at the stagnation line,  temperatures for its lateral surfaces where flow is still attached,  and temperatures for leeward separated wake zone surfaces.  These were figured for thermal re-radiation exactly balancing convective-only input,  as figured for a “dark” highly-emissive surface,  of thermal emissivity 0.8 as representative. 

Note that with the exception of only the longest-range suborbital mission,  all the rest of these data are under 1600 F,  and would permit exposed-metal construction of 316L stainless steel,  or of Inconel X-750,  or something in that same class!  And that even includes the return from the orbital mission!  Because of the stagnation zone temperature approaching 2000 F on the longest-range suborbital mission,  there needs to be some minimal heat protection in and near the stagnation zone.

In Figure 10 just below,  the format for surface pressures is similar:  trends of suborbital surface pressure vs entry speed at stagnation,  at lateral sides,  and in the separated leeside wake.  The orbital data again fall far off the trend lines.  There is a numbered key to relate missions to individual data points.  Note that no mission,  not even the longest-range suborbital,  exceeds 0.19 atmosphere anywhere.  That would be about 2.79 psi,  very modest indeed. 

Given the hot material strengths reported as part of Figure 9,  that means even a fragile extreme-low-density alumino-silicate ceramic composite could serve as heat protection.  So could ceramic fabric blanket or quilt-type materials,  if they survive wind shear.   Even a thin sheet of 2000 F-capable metal overlying mineral wool insulation would serve,  mounted only locally near the stagnation line. 

Conclusion:  the “hopper” could easily be designed to also serve as a low orbit taxi!

Figure 9 – Trends of Surface Temperatures vs Entry Speed  

Figure 10 – Trends of Surface Pressures vs Entry Speed  

Update 11-22-2023:  The following Figure 11 illustrates exactly why the surface emissivity must be high (a very dark or black surface color,  with a dull finish).  There are exposed metallic construction solutions and a refractory solution with simple alumino-silicate ceramics,  especially away from the stagnation zone,  if emissivity is high.  There ae only ablative solutions available if emissivity is low.

Figure 11 --  More Detailed Hopper/Taxi Heating Data

Surface Freight Transport on Mars

I have seen many notions for surface transport discussed on the New Mars forums,  mostly in the planetary transportation topic.  Nearly all of these suffer from the very serious downside of requiring the emplacement of significant infrastructure on Mars,  something very effort-intensive and very expensive,  here on Earth.  With transport costs to Mars being “astronomical” for the foreseeable future,  that is a fatal requirement for anything we might consider on Mars.

What one really needs to do first is determine a very fundamental characteristic:  is the cargo time-sensitive,  or is it not?  Transport of people over long distances is a very time-sensitive problem,  for which the solution is some sort of flight,  as Figure 1 suggests.  Most bulk freight is not time-sensitive at all,  and can travel very slowly,  similar to scheduled rail freight here on Earth.  Considering the expensive infrastructure issue,  what we need is a train that does not need any tracks. 

That small portion of freight to be transported that is time-sensitive is likely to be medicines and similar:  small packages that can fly with the people.

Figure 1 – Time-Insensitive Freight Should Go Slow,  On The Surface

The critical thing here is to minimize the infrastructure we must emplace to get any of these transportation ideas going.  The more you have to build,  the more it is going to cost,  for both the efforts,  and the materials,  plus the equipment with which to build it.  That’s true here on Earth,  which is why many of these ideas have never really been widely built,  even here.  On Mars,  it will cost a lot more still,  because of the interplanetary transport costs,  plus the development costs of re-making the materials and processes into those that will actually work on Mars.  As indicated in Figure 2,  about the only thing we have available that would actually work on Mars,  too,  would be graded dirt roads,  with “truck trains” moving on them.  True rail would cost less to run,  being lower friction,  but we have no way to manufacture steels on Mars,  including the extreme-cold adapted steels necessary there,  and no way to make the ties,  whether from steel,  concrete,  or wood.

Figure 2 – The Real Requirement is That There Be Almost No Infrastructure to Emplace

So,  it is the slow-moving “truck train” hauling the time-insensitive freight on graded dirt roads that we need for Mars.  This is really just a tractor pulling a string of freight wagons,  but internal combustion engines as we know them here on Earth are “right out”.  You not only have to carry the fuel,  but also the oxygen with which to burn it,  which considerably outweighs the fuel.  And unless you want to completely redesign the engines to handle 500-1000 C higher flame temperatures,  you will also have to carry diluent gas,  which in turn far outweighs the oxygen.  So,  in any practical sense,  you are looking at electric-powered tractors.

These can be solar electric powered,  however.  On a sunny day here on Earth at lower latitudes,  there is roughly a horsepower’s worth of energy per square yard falling on the ground.  At Mars,  there is about half that.  Call it 300 W per square meter as a rule-of-thumb.  Solar electric conversion efficiencies are around 20%,  so about 60 W per square meter of collector surface could be had,  for much of the day.  The trick is then to fit the freight wagons with solar panels on their roofs,  all connected electrically to power the tractor.  You just go very slow to stay within the power your solar collectors can generate.  But,  most bulk freight is very time-insensitive,  so who cares?

This thing doesn’t need a crew,  it can be self-driving between the spoil berms created from grading the road.  If programmed with a keep-right feature,  traffic on these roads can be two-way.  The basic characteristics and features are summarized in Figures 3 and 4. 

Figure 3 – The “Truck Train” Concept

Figure 4 – This Is How It Is Controlled

Now we need to verify feasibility with some numbers.  Trucks and train cars here on Earth are about 1/3 structure and 2/3 payload,  sometimes ¼-3/4.  Call it 30-70,  as a rule of thumb.  For a 100 metric ton loaded freight wagon,  you are looking at 70 tons of payload,  and 30 tons of chassis,  wheels,  drawbars,  containing-structure for the freight,  and solar panels on the roof. 

These things could use the very same giant rubber tires we use for off-road construction and mining work here on Earth,  but we may need to heat them slightly,  to prevent their cracking in the cold on Mars.  That can be done.  Tires on relatively smooth,  firm dirt have crudely 10 times the friction coefficient of tires on a paved road.  Using the Mars weight of a 100-ton loaded freight wagon,  augmented for climbing a 10% grade,  the drawbar power (drag x speed) is 20 KW at 0.33 m/s,  and quite a bit higher as you go faster.  There’s room for at least around 300-ish square meters of collectors atop each wagon.  Which in turn is why you select the lower of the speeds given in Figure 5.  

Figure 5 – This Is Why It Works (By the Numbers)

You can pull many freight wagons with a big powerful tractor.  Numbers are given in the figure for 10 and 100 wagons.  Power at the tractor is between 0.2 to 2 MW for these numbers.  That’s about like the power of a big mine loader as currently built here on Earth (somewhere near 1000 HP).  Which in turn means the same electric tractor that can pull the “truck train” can also be the road grader! You just power it differently.

The kind of thing I have in mind as the basis for designing the Mars tractor is pictured in Figure 6.  That is a big mine loader vehicle.  These are very tough,  very powerful machines.  That is the very thing we need for doing road grading,  and pulling heavy trains of freight wagons.  We just have to make it work on Mars,  in the cold and the near-vacuum.  That is why electric propulsion is preferred,  and a pressurized operator’s cabin is required. 

Figure 6 – The Mars Tractor Is a Known Revision of Something Like This That We Already Build

The basic notion is illustrated in Figure 7.  Something similar to a big mine loader is the design basis.  You remove the dump hopper and replace it with a battery bank and solar panels.  You remove the diesel power plant and replace it with the appropriate electric drive motors and gearing.   You replace the operator cab with a pressurized operator’s cabin for use on Mars. 

Rigged with a blade,  and powered by a nuclear generator aboard a shielded trailer,  you operate it manned for grading the road.  Unmanned and self-driving,  you pull the “truck train” with it,  powered by the solar panels on the freight wagons.

It will have to be manned for grading the road.  Even here on Earth,  “dirt work” is something that has so far proven to be not-automatable,  or it already would have been automated.  That’s slow multi-pass work,  with the operator’s judgement and skill showing up,  as knowing when the work is done “right” by the appearance of what he has done.  This aspect will be the same on Mars as it is here.

It does not have to be manned to pull the “truck train” on the finished roadway.  You remove the blade,  and delete the nuclear power trailer.  However,  I would leave the pressurized cabin in place and rigged with supplies,  for the off-chance unforeseen event that would require a human driver.  

Figure 7 – Converting a Mine Loader Design into a Mars Tractor Design That Does Two Jobs

The only real trouble is shipping such large pieces of equipment to Mars.  These will have to be shipped as individual components,  and assembled there on the planet Mars. That will be true until some very large interplanetary vehicles indeed,  eventually become available.  But it can be done,  with what we know and have available,  right now!