2020 Starship/Superheavy Estimates for Mars

Update 7-2-2020:  I found a mistake in my performance-estimating spreadsheet.  I used the unfactored delta-vee requirements for figuring mass ratios,  instead of the factored delta-vee requirements.  I corrected this in the spreadsheet,  and used those revised images to correct the spreadsheet results depicted here as Figures 5,  7,  and 9.  Then I corrected the results shown in Figure 10 by re-editing that figure. Where numbers changed in the text of the article,  I edited them with line-out and red text.

-----------------------------

These results are based upon the earlier article posted here,  titled “2020 Reverse-Engineering Estimates for Starship/Superheavy”,  dated 25 May 2020,  and also upon the earlier article titled “Interplanetary Trajectories and Requirements”,  dated 21 November 2019. 

The first examined the latest data available for the Starship/Superheavy designs,  and projected its performance into low Earth orbit (LEO),  complete with a look at what kind of tanker flights might be required to leave LEO.  That same set of weight statements and engine performances is assumed here.

The second examined trajectories to Mars,  including both Hohmann min-energy,  and faster trajectories,  and including direct entries,  and entries into low Mars orbit,  and to Phobos.  Of most interest here was the faster trajectory of an exactly 2-year period,  which provides a 2-year abort path if the direct landing on Mars cannot be made,  for any reason whatsoever. 

This article also examines direct entry and landing upon Mars,  starting from LEO,  via either Hohmann or the faster 2-year abort trajectory.  Further,  it looks at stopping in low Mars orbit (LMO),  followed by deorbit and landing on Mars.  It does not examine going to Phobos,  which would be broadly similar to stopping in LMO,  except a bit worse,  energetically.

Returns are launched directly from the Martian surface,  assuming a full refill using propellants made locally on Mars.  Whether one stops in LMO or just escapes directly,  the total departure velocity requirement is the same,  for the same return trajectory.  Stopping in LMO relieves a tight launch window from the surface,  replacing it with a fairly tight window for trans-Earth injection from LMO,  which is actually an easier target to hit,  in actual practice.

A summary illustration (not to scale) of the two basic transfer trajectories,  with the LMO option,  is given in Figure 1.  All figures are at the end of this article.

It must be understood at the outset that the kinematic velocity requirements for this-or-that trajectory are not what you design your vehicle with,  using the rocket equation!  Those velocity requirements must be factored-up appropriately to “mass ratio-effective velocity requirements”,  in order to size a credible design,  or make credible performance calculations.  That is because the rocket equation was mathematically derived for zero-gravity,  zero-drag conditions. The real world often is not.  See Figure 2.

I have been to the Spacex website many times over the last several years.  Among the other things posted there have been atmospheric entry simulations for their Starship vehicle.  There are two such:  one for entry at Earth,  and one for entry at Mars.  These are rarely posted at the same time,  and are quite different,  because the Martian atmosphere resembles the Earthly atmosphere above 105,000 feet (about 33 km).  See Figure 3. 

The Earthly entry trajectory ends hypersonics at about Mach 3,  above 40 km,  followed by a bending downward while pitching upward to dead-broadside descent angle-of-attack.  This dead-broadside fall is termed the “belly-flop” or “skydiver” maneuver.  It is supersonic up in the thin air,  and very subsonic down in the thick air near the surface.  Spacex lists this as 68 meters/second under about 5 km.  There, they pitch to tail-first while igniting the engines to land.  That is a very modest delta-vee on the order of 0.070 km/s,  but it has to be factored-up significantly to cover hover and divert allowances.

The Mars entry ends hypersonics at about Mach 3 very close to the surface:  about 5 km altitude.  Spacex’s simulation shows a lifting pull-up to decrease speed while gaining altitude,  to around Mach 2-ish at 10 km/s.  Right after that is when they pitch to tail-first and fire-up the engines to brake for landing,  at something like 900 meters/sec = 0.9 km/s.  Again,  that needs a significant factoring-up to cover hover and divert needs.

It may well prove to be,  that Spacex will need to fire up the engines earlier at Mars entry,  in order to augment lift with thrust to achieve the pull-up maneuver.  If so,  the landing burn delta-vee is even higher,  but I did not assume that to be the case for this analysis!

The Hohmann min-energy trajectory is depicted in Figure 4.  This is an ellipse with its perihelion at Earth’s orbit about the sun,  and its aphelion at Mars’s orbit about the sun.  Half of it covers the voyage to Mars,  the other half covers the return voyage.  For purposes of this analysis,  I only looked at Earth and Mars at their average distances from the sun.  Those variations do affect the results,  Mars more than Earth because of its more eccentric orbit.  But,  the averages are close enough to find out “what ballpark you are playing in”,  which is the point here.

For the baseline outbound voyage,  entry at Mars is direct from the interplanetary trajectory.  There is only a course correction budget needed along the way.  The three burns are LEO departure,  course correction,  and direct landing upon Mars.  The value of the course correction budget (0.5 km/s) is nothing but a guess,  but it is a good ballpark guess,  being about 2-3% of the typical velocity with respect to (wrt) the sun during the voyage.  Note that in this baseline there is no possibility of aborting the direct landing for reasons of adverse conditions at Mars.

For the baseline return voyage,  entry at Earth is direct from the interplanetary trajectory.  The same 0.5 km/s course correction budget is presumed for this leg.  The three burns are Mars departure,  course correction,  and direct Earth landing.  There is no possibility of aborting the direct Earth landing.  The difference is Mars departure,  which could be broken-up into two burns:  one to achieve LMO,  the other to escape onto the return interplanetary trajectory.  The total is the same,  either way.

What I got for this mission is the spreadsheet image shown in Figure 5.  The weight statements,  engine selections for each burn,  and engine performances,  are all listed in the spreadsheet image.  It covers both the voyage to Mars,  and the return to Earth,  as separate sections,  with different weights (payload values) for each leg.  Payload both ways exceeds what the “2020 Reverse-Engineering Estimates” article found to be the maximum payload deliverable to LEO.  The payloads shown are the max values to the nearest metric ton,  that gave positive fractional-ton remaining propellant,  after all burns were fully expended (the spreadsheet really is iterative).

Therefore,  to meet these weight statements,  not only must the Starship’s propellant be refilled to 100% of capacity on LEO,  but also some extra payload mass must be brought up to LEO and loaded on board.  Those means are not explored here.  However,  284 220 metric tons of total payload (people plus cargo) is a generous allotment that this design can deliver to Mars,  if an 8.62 month one-way voyage is acceptable. 

The Hohmann return payload of 153 144 metric tons is less,  because the net total delta-vee requirement to return is higher.  This mission spends about 13 months at Mars waiting for the orbits to be “right” to come home,  so you have 13 months to make 1200 tons of propellant!  The next opportunity is 26 months after that. The min round trip total is about 30.2 months.

The faster transfer trajectory that I chose to analyze is the 2-year abort return ellipse.  Its perihelion is at Earth’s orbit,  but its aphelion is far beyond Mars,  nearly to the main asteroid belt!  Its period is exactly two years,  so that if you follow it,  the Earth will be at its perihelion just as you arrive there.  In order to have Earth there when you are there,  the period of any such orbit must be an integer multiple of 1 year.  It is simply not a free-return abort orbit,  if the Earth is not there when you get there.  See Figure 6.

I used the same landing allowances and course correction allowances for the faster transfer case.  The LEO departure and Mars surface departure velocities are significantly higher,  because the energy of this orbit is also significantly higher.  That raises required mass ratio,  and lowers max payload.  Again,  the payloads I found were maximum to the nearest ton,  for a positive fractional-ton remaining propellant,  after all burns were fully expended.  What I found is the spreadsheet image in Figure 7. 

Max payload to Mars still exceeds equals what can be delivered to LEO without any separate freight deliveries!  Max payload returnable to the Earth is actually quite limited at only 39 34 tons!  This mission has about 15-1/2 months on Mars in which to manufacture the return propellant.  It is 2 years long overall.

As for stopping in LMO instead of direct Mars entry,  I only looked at the outbound voyage to Mars.  I did look at both the Hohmann min-energy transfer,  and the faster transfer using the 2-year abort orbit.  This added more burns to be made.  Now,  there are LEO departure and course correction as before,  plus an entry into LMO,  plus a deorbit from LMO,  and the same landing burn as before.  The faster transfer has the higher LEO departure and LMO entry requirements.  The final landing burn and the deorbit and course correction burns,  are all the same.  This is shown in Figure 8.

What I found for this,  is the spreadsheet image given in Figure 9.  I was quite surprised to find that it is indeed possible for a Starship to enter LMO and still have the propellant remaining to land.  It can only do this from the Hohmann min-energy transfer,  and only with very drastically-reduced payload (only 39 34 metric tons).  This is just not a possible landing abort mode,  unless you deliberately use Starship to transfer very small payloads to Mars.  From the faster transfer trajectory,  we are 21.5 33.6 tons short of the min landing propellant needed to avoid a fatal crash,  even at zero (!!!) payload.  That is just not feasible.

I have summarized all these results in Figure 10.  Using the “2020 reverse-engineering” numbers,  payload transferable to Mars looks to be more,  or at least equal to,  than what can be transferred to LEO,  for both Hohmann min-energy transfer (284 220 metric tons for 8.62 months one-way),  and for fast transfer on the 2-year abort orbit (197 149 metric tons for 4.26 months one-way).  The earlier article indicated a max of 149 tons could reach circular LEO,  with full recovery of both stages.

I looked at the LEO tanker issue in the “2020 Reverse Engineering” article.  Flying a cargo/passenger Starship at zero payload results in 133 metric tons of deliverable and transferable propellant,  to LEO.  That’s 9 flights to fully refill the one vehicle going to Mars,  assuming it arrives in LEO with sufficient propellant remaining to deorbit and land.  Flying a reconfigured Starship with extra tanks,  as a dedicated tanker design,  can deliver the full 149 metric tons to LEO.  That’s 8 flights to fully refill the one vehicle going to Mars.  Sorry,  8 or 9,  not 4 to 6.  The numbers do not support the optimism!

None of these tankers has the payload capacity to deliver what would increase the payload of the vehicle going to Mars from the minimum 149 tons.  Such increases would have to come from separate “freighter” flights.  There has been no way identified yet,  for transferring cargo from one Starship to another while in LEO.   No one has even mentioned this.

It is possible to stop in LMO before landing if you sharply reduce your payload before you ever depart LEO,  if you are on a Hohmann min-energy trajectory (78 50 metric tons for 8.62 months one-way).  But it is impossible,  even at zero payload,  to land without crashing,  if you fly faster!  The sharply restricted payload (78 50 versus 149 metric tons),  makes this option very undesirable,  but discarding it also eliminates a possible landing abort method,  if landing conditions unexpectedly prove adverse at Mars,  such as a giant dust storm with high winds.

Going to Phobos will not help this LMO-impracticality picture.   That slightly reduces the delta-vee requirements to rendezvous with,  and then land on,  the moon,  but it significantly increases the delta-vee requirements needed to reach the surface of Mars.  This has already been thoroughly explored in the article titled “Interplanetary Trajectories and Requirements“,  dated 21 November 2019,  on this site.  Other work I have done and not posted verifies that the least-energy path from Phobos to Mars is a burn to escape from Phobos,  a burn onto a transfer ellipse to LMO altitude,  a circularization burn at LMO altitude,  followed by a deorbit burn,  entry,  and a landing burn.

There are two critical items necessary to make this Starship/Superheavy design concept more practical:  (1) the ability to put closer to 300 220 metric tons of payload in LEO while maintaining an abort deorbit and landing capability,  and (2) coming up with a real tanker design.  Those are in addition to the rough-field landing capability problem,  and the inert mass growth problem that always occurs in experimental flight testing.  You have to increase the size of SuperHeavy to do that LEO payload increase,  so as to increase the staging velocity,  while still maintaining flyback capability for recovery. 

 Figure 1 – Types of Missions Considered

 Figure 2 – Factors Applied to Kinematic Delta-Vees to Obtain Mass Ratio-Effective Delta-Vees

Figure 3 – Trajectory Characteristics,  Final Speeds,  and Factors for Entry,  Descent,  and Landing

 Figure 4 – Characteristics and Velocity Requirements for Hohmann Min-Energy Transfer


 Figure 5 – Spreadsheet Image of Results for Hohmann Min-Energy Transfer AS EDITED 7-2-20

 Figure 6 – Characteristics and Velocity Requirements for Fast Transfer with 2-Year Abort

 Figure 7 – Spreadsheet Image of Results for Fast Transfer with 2-Year Abort AS EDITED 7-2-20

 Figure 8 – Hohmann and Fast Transfers to Mars Stopping in Low Mars Orbit Before Landing

 Figure 9 – Spreadsheet Image of Results for Stopping in Low Mars Orbit Before Landing AS EDITED 7-2-20

Figure 10 – Summary of Results for Transfers To and From Mars AS EDITED 7-2-20

Repeat-Pass Aerobraking at Mars?

This is a topic that often comes up,  as a potential way to save propellant for arrival into Mars orbit from the interplanetary trajectories.  There are two problems:  (1) Mars’s atmosphere is very thin,  so that the densities effective for drag deceleration at orbital speeds are quite close to the surface,  and (2) trusted sources such as the Justus and Braun tome on EDL (Entry,  Descent,  and Landing) indicate that the density of Mars’s atmosphere at braking altitudes is erratically variable by a factor of 2,  totally unlike the case at Earth.

Discussion of the Problems

The first problem brings up the risk of accidentally impacting the surface of the planet,  since that surface is so closely adjacent to any practical aerobraking trajectories.  Very precise trajectory control is required to make this work,  much more so than here at Earth.  It is not impossible,  but it is very demanding,  when this technique is to be used at Mars,  which lacks the navigation satellites we have here at Earth.

The second problem has two impacts:  (1) factor two density variation is factor 2 drag variation,  all else being equal,  which may be too large to assure aerobraking into orbit,  and (2) since the variation is erratic,  you cannot know the density that exists at aerobraking altitude until you are actually in the atmosphere attempting your aerobraking.  It will show up in your reduced speed and the peak gees decelerating to it.

Conditions of the Approximate Analysis

The altitude for “entry interface” at Mars is generally considered to be 140 km,  per Justus and Braun,  who recommend using 135 km for Earth entry interface altitude.   On Earth,  peak decelerations are up nearer 80-90 km,  while these occur lower on Mars,  primarily because Martian surface densities look like Earth densities at 30-35 km altitude.  I picked an arbitrary 50 km for my ellipse periapsis altitude,  to fall approximately in the max braking range on Mars. 

For speed at entry interface,  I used the 7.5 km/s that Spacex quotes for its peak entry speed with the Starship design.  That is intended for direct entry from the interplanetary trajectory.  Here I am investigating a braking alternative from those same conditions.  That entry speed corresponds to a velocity “far” from Mars near 5.6 km/s relative to Mars,  which implies a trajectory somewhat faster than a min energy Hohmann transfer orbit.

The remaining variable is the initial braking ellipse orbit apoapsis distance.  To be a ellipse at all,  this must fall between circular at the braking altitude,  and very eccentric with periapsis velocity very close to Mars escape.  I used a 10 m/s difference between escape and periapsis velocity as my largely-arbitrary criterion for the most elongated ellipse that I would consider “stable”.

Anything falling outside those limits cannot lead to repeat-pass aerobraking.  If the density and drag are too high,  the braking pass will inherently become an unplanned direct entry.   If the density and drag are too low,  braked speed exceeds escape,  and the vehicle bounces off the atmosphere into deep space.  Both outcomes are likely fatal,  especially bouncing off into deep space.

How the Repeat-Pass Aerobraking Works

The way this works is the vehicle approaches on a hyperbolic path to pass past Mars at the braking altitude.  As it approaches Mars,  gravitational pull increases its speed to the entry interface value,  approximated as:  Vint² = Vinf² + Vesc²,  where Vesc is computed at the braking pass altitude.  Vint is the interface speed,  and Vinf is the speed relative to Mars “far” from Mars.  For this analysis,  I just set Vinf such that I got Vint = 7.5 km/s.

During the pass,  density and velocity vary,  so drag varies.  There is a drag work integral with path length along the braking pass path.  That integrated drag work subtracts from the vehicle kinetic energy at entry interface,  so that kinetic energy (and vehicle velocity) are lower at the exit from the braking pass.  That final velocity must fall in the range of feasible periapsis velocities for the initial ellipse,  or else you enter direct if too low,  or bounce off into deep space if too high.

Once on a feasible ellipse,  there is drag braking at each periapsis pass,  which reduces vehicle mechanical energy and apoapsis altitude.  After multiple braking passes,  the orbit circularizes at the braking altitude,  and the vehicle then deorbits.  That is the concept behind repeat pass aerobraking.   See Figure 1.

Figure 1 – Repeat-Pass Aerobraking Conditions and Analysis Results

How the Analysis Was Conducted

I did not attempt to evaluate the actual drag work integrals.  Instead,  I looked at kinetic energy differences between entry interface speed and periapsis speed conditions for the range of feasible braking ellipses.  Those kinetic energy differences are the required drag work integrals. 

The max kinetic energy reduction (drag work integral) is for a circular orbit (degenerate ellipse of eccentricity zero) at braking altitude.  That orbit has about a 1.7 hour period.  The min credible kinetic energy reduction (drag work integral) is for an extremely-elongated ellipse with an apoapsis radius near 850,000 km.  Its period is nearly 3.3 months.  The ratio of those kinetic energy reductions (ratio of the drag work integrals) is about 1.39.  

Effects of Density Variability

For the same braking altitude (and it really won’t be,  but we are ignoring that effect here),  the drag work integral could vary max-to-min by a factor of 2,  simply because the density can vary by that factor of 2.  It varies erratically,  and you cannot know what that density really is,  until you have completed the braking pass with the wrong exit velocity,  which is far too late to adjust your braking pass altitude.  The factor 2 variation easily encompasses both unplanned direct entry and bouncing off into deep space.  Therefore you must be prepared to compensate “on-the-fly” as you exit the first braking pass by firing rockets to adjust your speed to the proper (or at least a feasible) value. 

It’s not that repeat-pass aerobraking won’t work,  because it most certainly will.  But at Mars,  the control of speed coming out of that initial braking pass requires very precise control and adjustment “on the fly”,  with a rocket burn of some significant size.  That plus the final circularization burn into a stable low Mars orbit will be less than,  but as much as some fair fraction of,  the direct orbit entry burn of almost 4 km/s.

Control is simpler,  and to more certain requirements,  with the orbit entry burn approach.  But if one has real-time velocity information,  and fast-response propulsion,  one can do a real-time-tailored delta-vee burn to exit the initial braking pass with the correct velocity.  You have to plan this for the very worst case,  even if you never use that propellant.

Conclusions

My conclusion is that the risk potential for a fatal outcome is just too high relative to the feasible braking orbit entry conditions,  unless you bring extra propellant and some very sophisticated,  precise,  and fast-acting controls.  This is not a technique that should be preferred for use at Mars!  One would have to add some serious rocket braking to prevent bouncing off,  or unintentionally landing direct but too steeply.  If you have to do that,  then you might as well just do rocket braking into orbit under much more certain control.

Details

The details of the calculations are shown in Figures 2 through 4.  These are spreadsheet images,  two images per figure.  I looked at a total of 6 ellipse conditions,  to see what the trend shapes looked like.  The approach to the bounce-off point,  just past the most elongated orbit,  is sort-of asymptotic in terms of kinetic energy,  so that energy estimate is insensitive to the exact max apoapsis criterion.

Remember,  because of unpredictable upper-atmosphere density variations,  the drag integral can vary through a factor of 2.  The ratio of drag integrals leading to a feasible ellipse is only 1.39.  Even if you do not care what ellipse you end up on,  the chances are very significant that you could see a fatal outcome (unintended direct entry or bouncing off into deep space),  without a significant propulsion allowance for adjusting speed as you leave the initial braking pass. 

If you care which ellipse you end up on,  you simply must have propulsive speed adjust capability coming out of that first braking pass. Either way,  the propulsion must be sized for the worst possible case,  which Murphy’s Law says will happen.

Figure 2 – Spreadsheet Images for Circular and Apoapsis at Phobos

Figure 3 – Spreadsheet Images for Apoapsis at Deimos and a One-Week-Period Orbit

Figure 4 – Spreadsheet Images for a One-Month-Period Orbit and the Max-Credible-Elongated Orbit

Thoughts On The Protests

Update 6-7-2020:  a version of this article just appeared Sunday 7 June in the Waco Tribune-Herald,  as a letter to the editor on the op-ed page.
------

Protest is as American as apple pie.  Looting and arson are not. 

Sometimes the distinction is blurred,  rendering this not a black-or-white,  this-or-that question.  Example: the Boston tea party also involved the destruction of private property.  It would be wise to remember that inconvenient little fact.

What is true today about these protests was also true in the 1960's and early 1970's.  There are (and were) the honest protesters,  there are (and were) the opportunists who want to use the chaos for criminal behavior,  and there are (and were) extremists (both left and right) who have their own agendas to take advantage of the chaos.

It is entirely too seductively easy to tar the honest protesters with rightful anger about criminal opportunists and violent political extremists. That mistaken identity and misplaced blame for the violence is usually what politically-motivated extremists want.  It would be well to remember that!

It is also entirely too easy for honest protesters to stray into wrongful behavior in the heat of the moment.  That last is the fundamental human characteristic that also makes lynch mobs so frequent in our history.  It would be well to remember that,  too.

Just food for thought,  guys,  from someone who saw this movie once before.

What we really need is twofold: 

(1) We need to address the underlying issue that is the reason for the current protests.  In this case,  that is structural racism still built into our society and our institutions,  up to and including lethal levels.  The anger being expressed about it is quite real,  and has been pent-up and festering since the Civil War.

(2) We need some way to distinguish in-the-field and real-time the honest protesters from the opportunistic criminals and the extremists promoting violence.  And,  we need a way to deal with the opportunistic criminals and violent extremists on-the-spot,  but without taking out inappropriate rage upon the honest protesters! 

I don't yet know how to do that second thing reliably.  But figure it out,  we must. 

The first thing requires that we examine our collective selves and root out the racism therein.  And quit teaching it to the next generations.  Simple,  but obviously very hard to do.  Or we would not still need to do this. 

On the Manned Spacex Launch

Update 6-6-2020:  a version of this article just appeared Saturday 6 June in the Waco Tribune-Herald,  as a board of contributors item on the op-ed page.
-----------------------------------

On Saturday May 30^th, I watched live TV coverage of the launch of two American astronauts toward the space station,  from Cape Canaveral.  It’s been almost a decade since I’ve seen that,  and that’s just too long.

This launch had no “window”,  it had to be on-time or not at all.  This effort was scrubbed for weather just last Wednesday,  because they could not wait out the bad weather.  The risk was a lightning strike on an ascending rocket and spacecraft,  something known not to be tolerable.

The launch and ascent looked “nominal” the whole way to orbit insertion,  where TV coverage ended.  They even successfully landed the booster first stage on the drone ship.  Spacex made this look easy and routine,  but,  believe me,  it is not.  Not yet.

Now,  Spacex has been launching unmanned,  robot-controlled cargo spacecraft to the space station for some years now.  And last year,  they successfully sent the manned version of this spacecraft to the station as an unmanned,  robot-controlled demonstration. 

This manned trip was the final demonstration.  From here,  Spacex should be cleared to send crews to the space station regularly for NASA.  And everything that flies is tested at the McGregor test site.  Tested by people who are paid well to do it,  and who spend money to live here.

Remember that when you hear that rocket testing noise:  continuous thunder is the sound of success.  It is the single ear-splitting “kaboom” that spells trouble!

NASA funded two contractors to do this astronaut ferry job to the space station.  Both need to be in operation,  in order to make this as reliable as possible.  The other contractor is Boeing.

Boeing still needs to do a successful unmanned,  robot-controlled demo before it can launch the final crewed demo.  They tried that a few months ago,  but it didn’t work right. 

Uncovering the failures to be fixed is the whole point of a testing program,  so having a problem with an unsuccessful flight is not a bad thing.  They will fix it and fly again.  They need our support,  too.

Now,  some of you may have heard about a Spacex rocket blowing up in south Texas about that same time.  That was a highly-experimental prototype for the new giant vehicle Spacex is trying to build.  This has nothing to do with the Falcon rocket and Dragon capsule that just launched with a crew.  It is a new,  future thing.

It blew up,  apparently from a propellant leak,  after a successful engine test.   It was supposed to do a gentle low-altitude “hop” sometime in the next few days,  but that won’t happen now,  because it was destroyed.  Such is the nature of early experimental testing of new designs.  Better to find the troubles early:  you lose fewer lives and less money that way.

This new giant vehicle is to be a huge transport to orbit,  first and foremost.  Refilled with propellant from tankers,  it can take large loads to the moon and Mars from there.  This is the wave of the future,  but it won’t happen,  without lots of experimental testing now. 

Which activity by its very nature is going to have some spectacular failures.  Like the one in south Texas. 

Why is this important?  First,  the space program has spin-offs that benefit the public.  It always has. 

Your Pyrex glass cookware resulted from warhead re-entry work in the 1950’s.  The desktop and laptop computers and the related cell phone technology that nearly all of you depend upon came from the computer rocket guidance systems developed in the 1960’s.  The weather predictions you depend on came from weather models and weather satellites developed to support spaceflight in the 1970’s.  And so on,  and so on.

Second,  some day,  if this gets inexpensive enough,  and reliably safe enough,  your kids may travel in space,  or from point-to-point on Earth through near space. 

In the 1960’s (but measured in today’s dollars),  the cost of sending a pound of payload to orbit was $10,000-100,000 per pound.  Because of Spacex and the rest of the satellite launch business entities,  that has been reduced to nearer $1500-2000/pound.  Bigger,  more reusable vehicles will reduce that further.

If it ever gets down under $100/pound,  that’s getting much closer to the price of a first-class airline ticket.  At which point the old dream of vacationing in an orbital hotel begins to become feasible.  We’re not there yet,  but the progress to date is absolutely astounding.

Third,  there is a need to protect the Earth from asteroid and comet impacts.  There is no better justification for both manned and unmanned space programs than this.  It requires sending both unmanned and manned craft out there to develop the protection means,  and then to use it when the need arises.  And it will arise,  at some time in the foreseeable future.

On the Pandemic

Update 6-20-2020:  A version of this article appeared in today's Waco "Trib" on the op-ed page.

Update 6-14-2020:  3 paragraphs in original article highlighted,  and two more appended below as part of this update,  which addresses the recent spike-up in cases as we re-open.

--------------------

When there is a new disease,  for which there are no treatments (other than support) and no preventatives (such as vaccines),  there is only quarantining,  at one level or another,  available to use as a tool to stop the spread of the disease.  This tool is centuries old,  but it works,  and combined with modern knowledge of the microbe theory of disease,  it can be rather effective. 

That’s the result of the accumulation of factual knowledge through science.  There are no “alternative facts”,  there are only lies.  It is now,  and it always has been,  best to use what we know to be true.    And there is no better way to know what is true than science!

When this SARS-CoV-2 (“Covid-19”) pandemic started,  we knew next-to-nothing about its characteristics.  Very quickly,  we learned that it is transmitted from person to person by virus particles contained in the tiny respiratory droplets we expel when we breathe,  talk,  cough,  or sneeze.  The virus particles cannot live in the air outside those droplets.  These droplets tend to fall to ground fairly rapidly,  and they can persist on surfaces where they fall,  or are otherwise deposited. 

That it spreads in droplets means that face masks are moderately effective in stopping disease transmission,  as the size of these droplets is large enough to be stopped by filter media through which we humans can still breathe.  That would include both paper and cloth masks,  even simple bandanas.  The physics of this says that in most situations,  the mask does not protect the wearer,  but it does protect others from any disease the wearer might have.  Only when up close,  really “in-your-face”,  does the mask significantly act to protect its wearer (such as medical workers).

That these droplets tend to rapidly fall to the ground is the origin of the 6-foot rule (or in other metric parts of the world the 2-meter rule) for physical distance between folks.  If you get closer,  that’s rapidly becomes the “in-your-face” situation.  Like the face mask,  this distancing is also moderately effective at stopping disease transmission.

Now,  acting together,  masks and the 6-foot rule are far more effective at preventing disease transmission than either is alone.  That’s why the CDC recommendations say do both.  It’s an “and” situation,  not an “either-or” situation.  And it is not supposed to be optional:  when I go to the grocery store,  I see too many shoppers not wearing masks,  and getting too close to each other.  Shame!

There is a third thing that can be done,  also moderately effective by itself.  That is wearing disposable gloves when handling items in public that are touched by many.  If you add that to the masks and the 6-foot rule,  there is little risk of transmission of disease,  such as when you go to vote. This would be true even if the crowds were large,  and the voting machines were not disinfected between voters (which they should be).  You will note in some news stories that workers in some businesses are using these three tools together.  Now you know why.

What we have learned over the last couple of months is that this SARS-CoV-2 / Covid-19 disease is more infectious than most ordinary influenzas,  and (more alarmingly) that a large fraction of infected people show no symptoms.  They do not know they are sick with it.  And for a couple of weeks while they have it,  they are spreading it like “Typhoid Mary”.  Not feeling ill is NO INDICATION you are not infected and infectious!  Again,  that’s just a scientific fact.  Doesn’t matter how you feel about it,  it simply “is”.

When this started,  we didn’t know so much.  It was wise to close down all gathering places of any kind where disease might be spread from person to person,  same as in centuries past. 

Now that we know as much as we do,  there is no reason these businesses and activities cannot resume,  as long as the transmission of disease is interrupted!  The main tools are face masks and the 6-foot rule,  augmented by gloves and face shields and frequent disinfection,  as necessary.  Every such activity is different. 

And it is morally incumbent on the operators of any activity to look closely at what is done,  by whom,  and how exactly it is done,  to determine how best to use these tools to stop disease transmission.  There are a few activities which cannot comply.  For public safety,  these must remain closed.

The ideology-based objections to these quarantining tools and the activity shutdowns have no basis in our law or our society’s history,  dating back to before the founding of the Republic.  Public health has always taken precedence,  even in George Washington’s revolutionary army.

There is a longstanding legal tradition that my right to swing my fist ends before I strike you with it.  Likewise,  your right to behave as you wish ends before you infect me by your behavior.  So,  wear the damned mask!  Stay 6 feet away!  Don’t open your business until and unless you can stop disease transmission in it.  Simple as that!

It could be a lot worse.  There are diseases (like chickenpox and measles) which can exist as bare virus particles floating around in the air.  No mask you could breathe through,  could ever stop them,  not by a long shot (science again).  Walk into any confined space where such are,  and you will be infected!  If this pandemic were one of those,  we could not safely reopen anything. 

It is not.  So,  count your blessings,  do this right,  and let’s get on with the war. 

And quit bitching about how public health orders violate your rights.  They don’t.  Your rights are NOT absolute.  They never have been.

----------------------------

Update 6-14-2020:

It’s official,  Covid-19 cases are on a sharp rise after re-opening,  even here in Texas.  Not enough people are taking the mask and 6-foot rules seriously.  Not enough reopened businesses have taken seriously giving their operations a hard look for how to prevent disease transmission.  This is exactly what the medical experts warned us would happen.  

To those who have not complied:  either you take this seriously,  or we will have to shut down and stay at home again.  If you take preventing disease transmission seriously,  then you don’t have to shut down or stay at home.  It’s your choices that make people sick and dead,  or not.    

Update 7-14-2020:  Well,  we opened up too quickly,  and with too many people behaving like there is no disease problem.  There is no reason why businesses and other things cannot be open,  as long as careful attention is paid to stopping disease transmission.  But too many people are either too stupid to understand the science,  or too incapable of critical thinking not to be deceived by internet lies and conspiracy theories.  The worst of which keep coming from our president and his ardent supporters,  instead of actual leadership in a crisis.  November is coming ---