Launch Costs Comparison 2018

Update 6-1-18:  see red text edits embedded below
------------------------------------------------------------
Update 6-16-18:  see blue text at end of article,  after the figures
------------------------------------------------------------
This article compares and correlates unit costs for launchers,  mostly those used commercially.  These data are based upon reported payload capacities and launch costs found in the literature.  The format is cost per unit delivered payload mass,  on the very important assumption that the launcher flies fully loaded.  All figures are at the end of this article.  Click on any figure to see any or all of them enlarged.  You can close that view box and be right back to viewing this article.  

Results are reported in millions of dollars per delivered metric ton,  and in dollars per delivered pound.  To estimate the unit cost when flying at less than full load,  simply divide these unit costs by the fraction of fully-loaded that you intend to fly. 

Scope includes the launchers used in the competitive satellite launch business,  plus a few launchers that were used,  but not competitively,  and the US Space Shuttle as representative of a large spaceplane.  Some of these launchers are no longer in service.  However,  the correlation results are used to predict unit costs for the NASA SLS block 1,  just for comparison.

Data,  Sources,  and Results for “Standard Low Earth Orbit”

“Standard Low Earth Orbit” is 23 degree inclination out of Cape Canaveral,  Florida,  to a 200 km orbit altitude.  This is what the reported payload delivery capabilities in the literature refer to.  These data are for one-way delivery of payload using a simple payload shroud,  not a recoverable capsule,  except for the Space Shuttle.  As researched,  those data are:

Of these,  the US Space Shuttle,  the Titan IVB,  and Falcon-1 are no longer in service.  There is a demonstrated history of reliability problems with Proton-M.  The Titan-IVB was retired in 2005.  The Falcon-1 retired no later than 2012.  The Space Shuttle retired in 2011.  The prices shown are for fully-expendable flights in the case of Falcon-9 and Falcon-Heavy.  There should be some small price break for flying reusably at reduced payload with these two launchers,  although how much is but speculation. 

The trends of unit cost per delivered metric ton (flying fully-loaded) are given in Figure 1.  Figure 2 shows unit cost per delivered pound.  Data are grouped and correlated as “competitive”,  “non-competitive”,  and “spaceplane.  The “competitive” launchers correlate flying fully loaded as:

                Unit cost $M/metric ton = 10.557 e^(-0.033 Wp)  where Wp = delivered payload, metric tons

                Unit cost $/delivered lb  =   4787 e^(-0.033 Wp)   where Wp = delivered payload, metric tons

The launchers marked “competitive” all compete in the commercial (and military) satellite launch businesses,  with market share in part depending upon price.  The launchers marked “non-competitive” never competed commercially,  and thus were not subjected to severe pressure on price.  The model for these assumes the same -0.033 Wp factor,  and uses a coefficient that forces the curve through the average of the Titan IVB and Delta IV Heavy data points: 

                Unit cost $M/metric ton  =   39.1 e^(-0.033 Wp)  where Wp = delivered payload,  metric tons

                Unit cost $/delivered lb = 17,720 e^(-0.033 Wp)  where Wp = delivered payload,  metric tons

This model was extended to 70 metric tons payload to estimate what should be expected for NASA’s SLS block 1 as shown in the figures ($3.878M/m.ton and $1759/lb).  That calculation corresponds to an expected launch cost of $271M,  when NASA’s actual launch cost estimate is $500M,  and its critics estimate twice that.  So instead of only 3 times more expensive than Falcon-Heavy (otherwise comparable in payload) as predicted by the correlation,  SLS block 1 is likely to be at least 6 times more expensive,  and it might even be 12 times more expensive.

Update 6-1-18:  the latest NASA estimates for launching SLS block 1 have grown to about ~$1-2B per launch.  Its critics put that closer to $4B.  For a 70 metric ton payload,  that's $14-56M/m.ton,  or $6500-$26,000/lb.  The range depends upon whose estimates you believe more.  That falls somewhere halfway between all the other launchers and the space shuttle.  SLS is well on its way to being the most costly launch vehicle in human history.  

The Space Shuttle (marked “spaceplane”) is quite different,  in that the delivered payload is but a small fraction of the mass of the recovered vehicle.  All the others are one-way trips to space,  with the delivered payload encased in a shroud.  There are no recovered capsules delivered by these launchers. 

The spaceplane model for the Space Shuttle assumes the same -0.033 Wp factor as the “competitive” launchers,  with a coefficient that puts the curve through the data point for the Shuttle:

                Unit cost $M/metric ton  =   131 e^(-0.033 Wp)  where Wp = delivered payload,  metric tons

                Unit cost $/delivered lb = 62,580 e^(-0.033 Wp)  where Wp = delivered payload,  metric tons

Re-Scaling Results for Delivery at the International Space Station (ISS)

I used the payload reduction fraction seen with the Space Shuttle as a constant applied to all the launchers still in service,  for estimating unit cost performance delivering to the ISS.  The ISS is located at a higher inclination and a higher orbit altitude.  For the same launcher technical performance,  a launcher’s max payload capability must be reduced when reaching for the more demanding destination. 

Flying with a 7 person crew,  the Space Shuttle is listed as 24 metric tons to standard low Earth orbit.  It could deliver as much as 27.5 tons,  but only with a smaller crew and less supplies.  Flying with a 7 man crew,  its capability to ISS is reduced to 16 tons.  That is 2/3 of the standard low Earth orbit capability with the same crew and supplies.  Applying this 2/3 factor “across the board” with the same launch prices produces Figures 3 (per ton) and 4 (per pound) below. 

I correlated unit cost estimates to ISS only for the “competitive” launchers that are still in service.  These are (of course) somewhat higher than for “standard low Earth orbit”,  because payload capability is lower,  while launch price is not.  This for one-way payload delivery using a simple payload shroud,  not a recoverable capsule.  That model is:

                Unit cost $M/metric ton = 15.428 e^(-0.048 Wp)  where Wp = delivered payload,  metric tons

                Unit cost $/delivered lb   =    6996 e^(-0.048 Wp)  where Wp = delivered payload,  metric tons

Estimating the Effects of Reusability

I based this estimate on what Falcon-9/Cargo Dragon has demonstrated to ISS with re-use of first stages,  when loaded to max cargo for ISS,  versus what I estimate the fully-expendable deliverable payload is to ISS.  The fully expendable estimate is 15.2 metric tons to ISS.  A fully-loaded (for ISS) Cargo Dragon is 8.8 metric tons.  That ratio is 0.5789,  and I assume it applies to Falcon-Heavy for its payload delivery to ISS with re-use of first stage cores.  The results are given in Figure 5,  for both full price and for an arbitrary modest price break:  80%-of-full-price,  representing savings from re-use. 

Estimating What SLS Block 1 Might Really Do (Standard Low Earth Orbit)

SLS Block 1 is said to deliver 70 metric tons to standard low Earth orbit.  NASA says it expects each launch to cost roughly $500M.  NASA’s critics say each launch might cost nearer $1000M = $1B.  Those data correspond to $7.14-to-14.28M/delivered metric ton or $3239-6478/delivered pound (flying fully loaded). 

The “non-competitive” launcher correlation predicts for SLS Block 1 a unit cost of $3.878/delivered metric ton or $1759/delivered pound (flying fully loaded).  Falcon-Heavy has an almost comparable payload (63.8 vs 70 metric tons),  with unit costs of $1.411M/delivered metric ton or $640/delivered pound (flying fully loaded and fully expendably).  SLS will never be reusable,  as that was never considered as a design requirement. 

SLS is expected to fly only once a year,  and not until 2019 or 2020.  Falcon-Heavy flew its maiden test flight in February 2018.  It is scheduled to fly at least two more times in 2018. 

Update 6-1-18:  as indicated above,  even NASA's own cost estimates for an SLS launch have grown dramatically,  as have the estimates of its critics.  It will come nowhere close to what the non-competitive launcher trend predicts,  being higher by a factor closer to 10 than 3.  This is only opinion,  but one would be tempted to say that this travesty is exactly what you should expect from corporate welfare mandated upon NASA by a Congress interested only in "pork".  

Other Launchers to Watch For (That Are Not Yet Flying)

There will be an Ariane 6.  Long March 5 may or may not be flying yet.  United Launch Alliance is designing a new heavy lifter to be called Vulcan.  The Jeff Bezos organization Blue Origin is designing a heavy lifter to be called New Glenn.  Spacex is working on a design called BFR which will be a super-heavy-lifter with a fly-back first stage combined with a second stage that is also a reusable spacecraft. 

Final Note:  Falcon-9 Cargo Dragon to ISS

Full price for a Falcon-9 launch is $62M.  This can send to ISS a Cargo Dragon totaling 8.8 metric tons.  Of that,  only 3.310 metric tons is actual deliverable cargo.  Using that 3.31 tons,  the effective unit costs for delivery to the ISS are: 

                $18.73M/delivered metric ton = $8495/delivered pound

Given the same 80% of full price with reusability,  as was used above,  these data reduce to:

                $14.98M/delivered metric ton = $6796/delivered pound

Compare those with what the Space Shuttle costs were,  delivering 16 metric tons to the ISS at $1.5B per launch:

                $93.75M/delivered metric ton = $42,517/delivered pound

These are the best guesses I have for Enhanced Cygnus on Atlas V 551,  and they are not accurate.  The max deliverable mass to ISS is 12.34 metric tons,  which has to be larger than the loaded Cygnus.  Data gleaned from multiple sites on the internet says the max payload to ISS inside the Cygnus is 3.5 metric tons max.  Cygnus cannot return to Earth.  Each launch is $153M.  Those unit costs are thus crudely:

                $46.M/delivered metric ton = $21,000/delivered pound

I have no reliable data on the cargo version of Soyuz,  riding the R-7 rocket.  Best guesses are max 2.4 metric tons of payload in the capsule,  and a launch cost on the order of $65M.  Those unit costs are:

                $27.1M/delivered metric ton = $12,300/delivered pound

Thus,  cargo Dragon on a Falcon-9 appears to be the most cost-effective means to deliver self-maneuvering and self-rendezvousing cargo to the ISS,  of all the vehicles that have done this task.  

Prior Similar Articles

This article replaces earlier postings on this site.  The best of the older postings is “Access to Space:  Commercial vs Government Rockets”,  dated August 7,  2015.  That one compares multiple rockets with the best inflation-corrected costs I could find or devise.  The one prior to that was “Revised Launch Cost Update” dated September 13,  2012.  It refers in turn to “Revised,  Expanded Launch Cost Data” dated May 26,  2012.  That one in turn was a revision to the original “Launch Cost Data” article dated January 9,  2012.  But this current posting is the best,  with the latest versions of the rockets,  and the most current costs I could find.  I did not inflation-correct costs from 2016 to 2018 values.

Figures Follow:

 Figure 1 – Unit Cost Comparison (per ton) to Standard Low Earth Orbit

 Figure 2 – Unit Cost Comparison (per pound) to Standard Low Earth Orbit

Figure 3 – Unit Cost Comparison (per ton) to ISS

 Figure 4 – Unit Cost Comparison (per pound) to ISS

Figure 5 – Unit Costs for Falcon Vehicles as Payload-in-Shroud to ISS with Re-Use

Update 6-16-18:

The competitive launcher curve-fit to LEO is $M/m.ton = 10.557 e^(-.033*Wp) where Wp is the metric tons of payload actually delivered from within a payload shroud.  It was derived for a variety of commercial launch vehicles,  including Spacex Falcon-9’s whose boosters were never re-flown more than twice,  if at all. Most were not reused.

From Spacex’s website,  posted prices and payloads-to-LEO for vehicles not flown reusably are $62M for a Falcon-9 carrying 22.8 metric tons,  and $90M for a Falcon-Heavy carrying 63.8 metric tons.  These correspond to curve-fit unit price predictions of 4.9748M/m.ton and 1.2858M/m.ton,  respectively. 

Using this and the payload,  back-predicted launch prices are $113M and $82M,  respectively,  compared to $62M and $90M as posted.  The discrepancy can be laid to the scatter in the data generating the curve fit.  It isn’t all that large at +26% for Falcon-9 and -9% for Falcon-Heavy.  But it is important to realize that predictions made this way could easily be wrong by a factor of 2. 

It is far too soon to be talking about what the launch price of a BFR/BFS ought to be.  However,  ignoring that,  and using this same correlation at the claimed 150 metric ton payload deliverable to LEO for the 2018 9-meter diameter version,  the predicted unit price is $0.07478M/m.ton,  and at that payload,  the back-estimated launch price should be near $11.2M.  That supposedly is the price Spacex would charge a customer to deliver his 150 metric tons of cargo to LEO.

I have estimated elsewhere on this site that this 9 meter diameter version of the BFS second stage requires 6 tankers to fully refuel its tanks with 1100 metric tons of propellant (80% liquid oxygen,  20% liquid methane),  enough for a one-way trip to a direct aerobraked Mars retropropulsive landing,  with 150 metric tons of payload.  That’s seven launches for each BFS that departs for Mars,  totaling $78.4M in direct launch costs for each BFS embarking on that trip.

Assume for the sake of argument the wild guess that ~$10B are left to spend making this design flight-ready for the Mars trip.  Assume also as another wild guess that another ~10B are needed to fully develop and field test-verify all the equipment needed to sustain a small crew on Mars for 2+ years,  and to emplace there the means to manufacture propellant locally for the return home (1100 metric tons per returning BFS).

Assume also that one ship gets lost and destroyed trying to make its landing.  Could easily be 2 or 3,  but assume it is only one.

Spacex’s posted plans call for 2 uncrewed BFS vehicles each carrying 150 tons of cargo to land during the first Mars opposition,  and 4 more the second opposition,  with two of those carrying small crews and lots of cargo totaling 150 metric tons.   That’s 6 successful BFS flights to Mars,  plus the one we assumed lost,  for a total of 7 attempts. 

At the estimated aggregated launch cost of $78.4M per BFS sent to Mars above,  that’s $549M in launch costs for the 7 attempts;  just under $0.6B.   Add in the $20B to prove out the vehicle and to prove out the hardware,  and you have a temporary base on Mars for under $21B!  Remarkable!  And note how launch costs are miniscule compared to the hardware development and verification costs that we have assumed.  We could easily double or triple the launch prices,  and not really affect this outcome.

Now,  the predicted $11.2M per BFR/BFS launch is for a matured,  well-proven design that flies routinely and reliably.  That’s what the curve-fit really represents.  Right now this thing exists only as a concept in presentations.  A whole lot of resources must be spent developing this vehicle and turning it into that near-perfectly-reliable transport.  That what the wild guess of ~$10B is all about.

The same applies to both crew life support technologies and most especially the on-Mars propellant production hardware.  If that doesn’t work,  the crew dies on Mars.  There is no rescue,  because without refilling on Mars,  the BFS is a one-way trip.  So the propellant production plant has to work!  And it has to produce some 1100 tons of propellant to refill that BFS which returns the crew,  before their supplies run out.  This is not a small piece of equipment!  That’s what the other wild guess of ~$10B is all about.

So,  no matter what,  you have ~$20B in vehicle and hardware development and prove-out costs to amortize across the flights this vehicle will make.   The more it gets used (for any missions,  not just Mars),  the lower that amortized impact on any one vehicle. 

If the only thing ever done with this vehicle is those 7 attempts at Mars,  then there must be a total of 49 BFR/BFS vehicles launched.  $20B spread over 49 vehicles is $408M per vehicle.  That needs to be added to the otherwise-mature launch price of $11.2M,  for about $419M per launch. 

If you do missions to the moon in a similar way to Mars,  that’s another 49 flights under the same assumptions otherwise,  for a total of 98 flights to amortize the development costs.  That would be $204M per vehicle,  for a total price of $215M per flight.

Now,  if the same vehicle gets routinely used to LEO,  that’s more vehicles over which to amortize the development costs.  Say,  100 orbital deliveries,  for a total of 198 flights.  That’s $101M amortized per vehicle,  for a suggested launch price of $112M per vehicle.  

This is the “mass production effect” in action.  Unless somebody else foots some or all of the development bills,  I fear the BFR/BFS vehicle launch price can never be as low as it could be.  But then again,  my curve fit may have been extrapolated too far.   We’ll just have to wait and see what really happens.  One thing is sure:  this thing will be way cheaper to use than NASA’s SLS ever could be!

Weather Versus Climate

This sketch explains why a cold winter does not disprove global warming,  despite what the skeptics so love to claim.

While only an over-simplified  sketch,  the picture speaks for itself.  If you see both harsher winter cold snaps and hotter summer heat waves,  that's one of the tell-tale symptoms of global warming!

As indicated in the sketch,  the extra heat energy goes into powering wider weather swings.

We can always debate as to why the warming might be occurring.  But it is definitely occurring,  despite the harsh winter conditions of January 2018.

If you want to see something about causality,  go see the article titled "Do We Fight Global Warming or Not?",  dated 4-15-17,  on this website.

Alabama Special Senate Election Outcome 2017

Well,  the voters of Alabama selected a Democrat rather than an alleged child molester to be their senator in the special election of 12-12-17.  That's a good thing,  but there's a downside.

They only voted that way by around a percent or so margin.  That means very nearly half the voters in Alabama that day actually preferred the child molester to represent them,  just for the political party advantage.

When the voters are so deluded by party propaganda as to effectively have no ethics,  then why is it a surprise that so many politicians are similarly detestable?

A Better Version of the MCP Space Suit?

This is a concept proposal for a better version of the mechanical counter-pressure (MCP) space suit.  It combines the best features and eliminates the worst disadvantages of the particular two MCP design approaches upon which it is based.  These are the “partial pressure” suit of the 1950’s and the “elastic space leotard” of Dr. Paul Webb.  The result should be a lightweight,  supple (non-restrictive) suit that with suitable unpressurized outerwear,  can be used on pretty much any planetary surface even if totally airless,  or even in space.  It need not use exotically-tailored materials in its construction.  It should be relatively easy to doff and don.

This article updates earlier articles on this subject.  Those are:

Date           title             

2-15-16     Suits and Atmospheres for Space  (supersedes those following)

1-15-16     Astronaut Facing Drowning Points Out Need for Better Space Suit

11-17-14    Space Suit and Habitat Atmospheres

2-11-14      On-Orbit Repair and Assembly Facility

1-21-11     Fundamental Design Criteria for Alternative Space Suit Approaches

The idea here is to combine the two demonstrated approaches that both apply the fundamental MCP principle:  the body needs pressure applied to its skin to counterbalance the necessary breathing gas pressure.  The body simply does not care whether this counter-pressure is applied as gas pressure inside a gas balloon suit,  or is exerted upon the skin by mechanical means.

The first article cited in the list above (“Suits and Atmospheres for Space” dated 2-15-16) determines that pure oxygen breathing gas pressures from 0.18 atm to 0.25+ atm should be feasible.  How that was calculated is not repeated here.  My preferred range of helmet oxygen pressures is 0.18 to 0.20 atm,  for which wet in-lung oxygen partial pressures range from 0.11 to 0.13 atm,  same as the wet in-lung oxygen partial pressures in Earth’s atmosphere at altitudes between 10,000 and 14,000 feet. 

However,  only 0.26 atm gives you the same wet in-lung oxygen pressure as sea level Earth air.  The 0.33 atm used by NASA is entirely unnecessary,  unless to help overcome the exhaustive efforts necessary to move or perform tasks,  in the extremely stiff and resistive,   heavy,  and bulky “gas balloon” suits they use.

The 1940’s design that operationally met the need for extreme altitude protection for short periods of time was the “partial pressure” suit of Figure 1,  in which compression was achieved with inflated “capstan tubes”.  These suits were widely used into the 1960’s.  The capstans pulled the non-stretchable fabric tight upon the torso and extremities.  This provided the counterpressure necessary for pressure-breathing oxygen during exposures to vacuum or near vacuum,  for durations up to about 10 minutes long.  This was for bailouts from above 70,000 feet,  and would have worked for similar short periods even in hard vacuum.  Hands and feet were left uncompressed,  but for only 10 minutes’ exposure,  these body parts could not begin to swell from vacuum effects. 

The advantages of this design were (1) ease of doff and don,  (2) it was simple enough to be quite reliable,  and (3) it was not very restrictive,  whether the capstan tubes were pressurized or not.  The disadvantages were the achievement of rather-uneven compression,  and leaving the hands and feet completely uncompressed.  This limited the allowable exposure time by (1) uncompressed small body parts begin swelling in about 30 minutes,  and (2) between the uncompressed parts and the uneven compression achieved on the extremities,  blood pooling into the under-compressed parts could lead to fainting within about 10 to 15 minutes. 

Figure 1 – Partial Pressure Suit Design Used From the late 1940’s to the Early 1960’s

In the late 1960’s,  Dr. Paul Webb performed striking experiments with an alternative way to achieve mechanical counterpressure upon the body.  He used multiple layers of elastic fabric (the then-new panty hose material) as a tight-fitting leotard-like garment.  This was not a single-piece garment.  It achieved more-uniform compression on the torso and extremities than did the older partial pressure suit.  Dr. Webb included elastic compression gloves and booties,  so that the entire body was compressed,  removing the time limits.  Breathing difficulties were solved with a tidal volume breathing bag enclosed by an inelastic jacket. 

Breathing gas was pure oxygen at 190 mm Hg pressure (0.25 atm) fed into the helmet from a small backpack with a liquid oxygen Dewar for makeup oxygen.  This type of garment was very unrestrictive of movement,  and was demonstrated quite adequate for near-vacuum exposures equivalent to 87,000 feet,  for durations up to 30 minutes.  It was intended for possible application as an Apollo moon suit,  but could not be made operationally ready in time.  It has been mostly forgotten ever since.

The advantages are very unrestricted movement,  very light weight (85 pounds for suit plus helmet plus oxygen backpack),  and no need for a cooling system:  you just sweat right through the porous garment,  same as ordinary street clothing.   Plus,  the garment’s pieces were quite launderable.  Dr. Webb’s test rig is shown in Figure 2.  6 or 7 layers of the panty hose material provided adequate counter-pressure.

Figure 2 – Dr. Webb’s “Elastic Leotard” MCP Space Suit Prototype as Demonstrated

The disadvantages were essentially just difficult (time-consuming) efforts to don and to doff the garment’s pieces,  precisely because they were inherently very tight-fitting.  For use on a planetary surface or out in space,  one treats the suit as “vacuum-protective underwear”,  and adds insulating or otherwise protective non-pressurized outerwear over it.  So protection from hazards is not a disadvantage at all,  but only if one uses the vacuum-protective underwear notion. 

The main advantage of Dr. Webb’s “elastic space leotard” over the “partial pressure” suit was the more even (and more complete) compression achievable with the elastic fabrics.  The main advantage of the “partial pressure” suit over the “elastic space leotard” was the ease of donning and doffing the garment,  when the capstan tubes were depressurized,  releasing the fabric tension.  Both approaches offer very significant advantages over the “gas balloon” suits in use since the 1960’s as space suits:  lighter,  launderable,  and far,  far more supple and non-restrictive for the wearer. 

That suggests combining both of the successful MCP design approaches (inflated capstans and elastic fabrics) into a single mechanical counterpressure suit design.  The capstans apply and relax the tension in the fabric which provides the counter-pressure on the body,  and the elastic fabric makes the achievable compression far more uniform.  What is required from a development standpoint is experimental determination of the number of layers of elastic fabric required for each piece of the garment,  in order to achieve the desired compression in every piece. 

If done this way,  there is no need for directionally-tailored stiffness properties in specialty fabrics,  the basis of Dr. Dava Newman’s work with mechanical compression suits (see Figure 3).   Ordinary commercial elastic fabrics and ordinary commercial joining techniques can be used.  In other words,  pretty much anyone can build one of these space suits!

Figure 3 – Dr. Dava Newman’s MCP Suit Based on Directionally-Tailored Fabric Properties

So,  the MCP suit proposed here has certain key features (see list below).  It will resemble the old “partial pressure” suits,  except that protective outerwear (insulated coveralls,  etc.) get worn over the compression suit itself,  and the helmet is likely a clear bubble for visibility.  There is an oxygen backpack with a radio.  There is no need for any sort of cooling system.  Everything is easily cleaned or laundered free of dust,  dirt,  sweat,  and similar contamination.

Key features list:

#1. Pressurized capstan tubes pull the elastic fabric tight whenever the helmet oxygen is “on”,  but depressurize and slack the garment tension when helmet oxygen is “off”.  The capstan tubes are just part of the oxygen pressure breathing system.  Slacking the fabric tension makes doff and don far easier.

#2. The multi-piece garment is composed of multiple layers of elastic fabric to provide the desired level of stiffness that will achieve the desired level of compression in each piece of the garment.  This depends upon both the shape of the piece,  and upon how much circumferential shortening is achieved by inflating the capstan.

#3. The pressure garment is vacuum-protective underwear,  over which whatever protective outerwear garments are worn that are appropriate to the task at hand.  For example,  the wearer might need white insulated coveralls and insulated hiking boots,  plus insulated gloves.  One could even add some sort of simple broad-brimmed hat to the helmet if sunlight were intense. 

#4. The clear bubble helmet is attached to the torso garment piece. This torso garment piece also incorporates an inelastic jacket surrounding a tidal volume breathing bag.  Helmet,  breathing bag,  and capstans all pressurize with oxygen from the supply simultaneously,  and are (in fact) connected.  All are activated by one on/off control.

#5.  The oxygen backpack is just that,  no cooling system required.  It probably uses liquid oxygen from a Dewar as make-up oxygen,  has regeneratable carbon dioxide absorption canisters,  and a battery-powered radio.   It might also contain a drinking water feed connected to the helmet.  Attitude and translation thrusters for free flight in space can be a separate chair-like unit,  and this function is entirely unnecessary on a planetary surface.

#6. For concave body surfaces and complex shapes like genitalia,  the pressure suit can incorporate semi-fluid gel packs that surround these body parts,  making the body effectively convex everywhere.

How all this works together is shown conceptually in Figure 4.

Figure 4 – How the Capstans and Elastic Fabric Work Together for an Improved MCP Suit

About the only caveat might be that the breathing gas pressure could be too small to also serve as the capstan inflation pressure.  If that should prove to be true,  then there need to be two final pressure regulators in the oxygen backpack,  instead of just one.  That problem can be easily solved!

Reverse-Engineering the ITS/Second Stage of the Spacex BFR/ITS System

Update 4-17-18:  since writing this article,  I have gone to the Spacex website,  where the 2017 presentation materials and more are posted about this design.  I have re-visited my reverse-engineering of the capabilities of this vehicle in greater detail with greater fidelity to reality,  in all its complexity.  I have posted that new,  improved analysis as "Reverse-Engineering the 2017 Version of the Spacex BFR", dated 4-17-2018. I do recommend that readers use that newer analysis article, rather than this one.

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

The “giant Mars rocket” proposed by Spacex has reduced in size somewhat since its first reveal at the Guadalajara meeting.  The term “BFR” is now beginning to refer to the first stage of the two-stage system,  which flies back and lands for reuse.  The term “ITS” more properly applies to the reusable second stage,  which apparently has two forms.  Those are the cargo/passenger craft that goes to destination after refilling on-orbit,  and a flyback tanker that provides the refill propellants on-orbit. 

Data published by Spacex at the latest meeting indicate a cargo/passenger vehicle that summarizes as given in Figure 1.  Grossly,  this is a 9 m diameter vehicle about 48 m long,  with a dry mass of about 85 metric tons,  and propellant tankage that holds about 240 metric tons of liquid methane and 860 metric tons of liquid oxygen.  Stated payload weights are 150 metric tons on ascent (and presumably to destination), and “typically” 50 metric tons on return.  Characteristics of the tanker form are less clear,  but it seemingly has a lighter dry weight of about 50 metric tons.

Figure 1 --  Estimated Characteristics of ITS Per 2017 Revelations

These two versions presumably share the same ascent propellant tankage and engine cluster.  Those engines include both sea level and vacuum expansion forms of the same Raptor engine,  with a nominal chamber pressure of 250 bar,  and deeply-throttleable to 20% thrust.  The cluster has 4 vacuum engines of 1900 kN thrust each at 375 sec vacuum specific impulse,  and two sea level engines of 1700 kN thrust each,  and specific impulses of 356 sec in vacuum and 330 sec at sea level.  Exit diameters are 1.3 m and 2.4 m for the sea level and vacuum forms,  respectively.  (I did not correct sea level thrust to vacuum.)

I am presuming here that second stage operation during launches to Earth orbit takes place in vacuum,  so I use the vacuum thrust data for both versions of the engine.  Each type’s thrust is therefore associated with a propellant flow rate via its specific impulse.  Summing these gets a total full thrust and a total propellant flow,  and thus an effective “average” vacuum specific impulse with all six engines running,  for an effective exhaust velocity of about 3.5762 km/sec.  That calculation summarizes as follows,  where effective cluster specific impulse is total thrust divided by total flow rate (Figure 2).

Now,  on the assumption that both forms of the vehicle have the same ascent propellant tanks and quantities (totaling 1100 metric tons of propellants),  the following weight statement and delta-vee table applies (Figure 3).  For the tanker,  the first-listed payload of 150 tons is assumed from the cargo passenger version.  The second is back-calculated from holding tanker delta-vee capability to be the same as the heavier ascent form of the cargo/passenger vehicle.   

To do that,  one finds the required mass ratio from the delta-vee,  then solves the mass ratio build-up for the unknown payload:

                Wpay = [Wp – (MR – 1)Wdry] / (MR – 1)

What I find very interesting here is that Spacex seems to have said it takes 6 tankers to fully refill an ITS on orbit for its voyage to destination.  If you look at the heavier tanker that gets the same 6.2 km/sec delta-vee as the fully-loaded cargo/passenger form,  then 1100 metric tons of propellant divided by an estimated 184.7 metric tons per tanker equals 5.956 (almost exactly 6) tankers required.  So the tanker at 50 tons dry weight seems to hold 1100 tons of ascent propellant,  and just about 185 more tons of propellant-as-payload with which to refill a cargo/passenger ITS on orbit.  It would appear this estimate is then just about right.  It does presume all 6 engines running all of the time.

Using BFR/ITR at Mars

For a trip to Mars from low Earth orbit,  the departure delta-vee for a Hohmann minimum-energy orbit to Mars is around 3.71 km/sec at average orbital conditions.  For a direct entry without stopping in Mars orbit,  you let the planet hit you from behind,  as the planet’s orbital velocity is faster than the transfer orbit’s aphelion speed.  Velocity at entry interface will fall in the 6 km/sec range,  and aerodynamic drag kills most of that to about 0.7 km/s coming out of hypersonics fairly deep in the Martian atmosphere.  Double or triple that for the landing burn:  about 1.5-to-2 km/sec delta-vee requirement. 

That’s crudely 5.21 to 5.71 km/sec delta-vee required to make a direct landing on Mars,  with just almost 6.2 km/sec available.  The difference can be used to fly a somewhat higher-energy transfer orbit,  for a shorter flight time than 8 months.  Faster is possible if payload is reduced.

To return,  the ITS is refilled with in-situ propellant production on Mars.  It will need around 6 km/sec delta-vee capability to launch and escape directly,  with enough energy to achieve the return transfer orbit.  We assume a direct entry at Earth,  which means in turn we run into the planet from behind,  since vehicle perihelion velocity is higher than Earth’s orbital velocity. 

It will be a very demanding entry interface speed (well above 11 km/sec):  this is what stresses the heat shield,  not entry at Mars.  But,  the vehicle will come out of hypersonics at about the same 0.7 km/sec moderately high in the atmosphere.  It will need at least 3 times that as the landing burn delta vee requirement,  because the altitude is higher,  and the gravity is stronger.  Call it 2 km/sec as a “nice round number” to assume.

The total delta-vee requirement to ascend from Mar’s surface and achieve a direct transfer orbit and a powered landing on Earth is therefore in the neighborhood of 8 km/sec.  That is just about what the ITS cargo/passenger version seems capable of,  if restricted to about 50 metric tons return payload.  Again,  that particular payload correspondence lends confidence to these otherwise-guessed numbers. 

It also points out how critical in-situ propellant production will be for using this vehicle on Mars.  Unless this vehicle is refilled locally with the full 1100 metric ton propellant load,  it is stranded there!  Each launch from Mars requires 240 metric tons of locally-produced liquid methane,  and 860 metric tons of locally-produced liquid oxygen.  Launch opportunities are 26 months apart.  Required production rates are thus 9.23 tons/month methane,  and 33.08 tons/month oxygen,  at a bare minimum,  per launch.

BFR/ITS For the Moon

Some have pointed out that this vehicle could also visit the moon.  To leave Earth orbit for the moon,  the delta-vee requirement about 3.29 km/sec.  The delta-vee to arrive into low lunar orbit is just about 0.8 km/sec,  or to land direct,  about 2.5 km/sec.  Those one-way totals are 4.09 km/sec to lunar orbit,  and 5.79 km/sec to land direct (remarkably close to the Mars value at min energy transfer). 

To return by a direct departure from the lunar surface requires about 2.5 km/s,  or from orbit about 0.8 km/sec.  Landing at Earth is largely by aerodynamic braking,  but requires about a 2 km/sec landing burn.  Therefore,  total delta-vee requirements to return are 4.5 km/sec from the surface,  or 2.8 km/sec from lunar orbit.

One could conclude that the ITS could ferry cargo to lunar orbit and return entirely unrefilled,  a trip requiring total 6.89 km/sec delta-vee capability.  This is not available at 150 metric tons of payload,  but it is available at something a little larger than 50 tons.  I get about 102 metric tons of payload. 

The requirements to land and return entirely unrefilled would be 10.29 km/sec,  which is out-of-reach even at only 50 tons payload.  To use the ITS on the lunar surface will require propellant production on the moon,  although likely at somewhat lower rates and quantities than at Mars.

Guessing Reusable Performance of BFR

A related point:  if we presume the fully-loaded ITS uses essentially all of its 1100 tons of propellant achieving low Earth orbit,  we can back-estimate the delta-vee that is actually available from its BFR first stage,  even allowing for flyback.  Earth orbit velocity is just about 8.0 km/sec.  Allowing 5-10% gravity and drag losses for a vertical ballistic trajectory,  the min total delta vee is about 8.4-8.8 km/sec.  About 6.1 of that is from the ITS second stage.  The first stage need only supply 2.3-2.7 km/sec,  which means the staging velocity is just exoatmospheric at around 2.5 km/sec.  It should easily be capable of ~5 km/sec,  so the difference is for flyback all the way to launch site,  and propulsive landing.

Suborbital Intercontinental Travel

Finally,  there has been some excited talk about using the BFR/ITS for suborbital high speed transportation across intercontinental ranges here on Earth.  That is a ballistic requirement similar to that of an ICBM.  The burnout velocity of the typical ICBM is around 6.7 km/s.  Allowing 5-10% margin for gravity and drag losses,  the delta-vee necessary to fly intercontinentally is 7 to 7.3 km/sec,  plus for the ITS,  about 2 km/sec for the landing burn.  Total is thus 9 to 9.3 km/sec delta-vee. 

This is way beyond the delta-vee capability of the ITS stage alone,  notwithstanding the fact that 4 of its 6 engines will not operate at sea level,  and even if they did,  total 6-engine thrust of the ITS stage (1100 kN) is less than its weight (1300 kN or more).  But this delta-vee is within reach of the two-stage BFR/ITS combination (6.2 to 7.9 km/sec ITS and ~2.5 km/sec BFR for 8.7 to 10.4 km/sec),  and likely with a little less payload than the 150 tons typical to Mars.  Maybe something in the vicinity of 100 tons. 

Final Remarks

These estimates are rough.  I did not correct sea level thrust to vacuum for one thing,  my delta vee requirements are approximate for another,  and I did not explore the effects of using only the vacuum engines for higher specific impulse out in space.  

Even so,  these results are very intriguing.  These calculations were made pencil-and-paper with a calculator.  Nothing sophisticated.  

ASUS Hardware, Windows Software? Never Again!

My ASUS X553M laptop with factory Windows 10 operating system is a low-quality,  unreliable piece of crap!  So is its operating system!  (Its predecessor was a Toshiba laptop running Windows 8/8.1.  The hardware failed at age 2:  the display hinges broke.  I hated Windows 8 from the moment I saw it.)

This ASUS machine/Windows software combination has several very serious issues that Best Buy’s Geek Squad cannot,  or will not,  help me with.  All these major issues are fatal,  as far as my estimate of quality is concerned.  That list follows below.

I would appreciate comments from readers as to what machines or operating systems might possibly be acceptable (since this machine and operating system are so very clearly not). 

I need to do word processing,  powerpoint-type slides,  spreadsheet work with plotting,  and a shell within which to run old-time DOS software.  I need something that can use wi-fi to access the internet and email.  I want a battery pack that I can pull,  to force a restart,  when all else fails.

ASUS X553M / Windows 10 Fatal Issues List:

#1. The screen dims and flashes or flickers,  when not plugged into the AC power supply.  This renders the machine unusable,  in spite of the battery being charged.  When the issue first started,  it did this with about 50% battery charge remaining,  as indicated on the display.  This rapidly got worse over a period of only months,  accelerated to starting the flicker at 90% battery indicated.  Now it will not run without flashing even at 100% indicated charge state.  Nothing in the Windows settings affects this. 

#2. The machine turns off its wi-fi device spontaneously,  without warning,  and for no perceptible reason.  This happens erratically and unpredictably.  The frequency with which it occurs is increasing as time goes by.  More of the time,  It still sees the wi-fi network,  and will reconnect if you command it.  But for a significant portion of the time,  it does not see the wi-fi network,  and so cannot be commanded to reconnect.  The only recourse in that case is reboot. 

#3. This machine on occasion locks up without warning,  rendering the keyboard and the mouse totally inoperative.  The only way to deal with this is a reboot.  It always loses all data up to the last save.

#4.  I cannot trust the reboot to be effective,  unless I unplug the AC power,  and either select full shutdown (not restart),  or else use the power switch.  I have noticed that the tiny indicator lights do not go out,  and that the issues the reboot was supposed to correct do not reliably get corrected,  unless I go for the complete shutdown with no AC connected.  There is no battery pack to pull,  as the battery is all-internal.  

#5.  The machine erratically and unpredictably ignores clicks of the mouse.  This problem comes and goes erratically. 

#6.  The keyboard has unreliable keys,  and a slow response to keystrokes.  You can type fast,  and it will miss a lot of letters.  Some are worse than others.  Those will often ignore slow repeated keystrokes,  even ignore continuous hold-down of the offending key.  Plus,  the symbols wore off the keys in only a year.

#7.  I haven’t seen a stable operating system out of Microsoft since DOS,  which would fit on a 1 megabyte floppy disk.  The entire fundamental Windows concept is flawed,  forcing people to learn a second language (icons),  which was (and still is) unnecessary.  The last DOS machine I had also had a little shell program (from a German company) that did a text-based point-and-click mouse controlled interface.  This interface did everything for file navigation that Windows ever did,  but would fit on another 1 megabyte floppy disk without even filling it. 

#8. Windows 8/8.1/10 are all useless pieces of crap totally bogged down with useless touch-screen crap that is totally inappropriate to an ordinary laptop.  That kind of marketing arrogance totally negates any possible past reputation Microsoft ever had for quality or for customer service. 

#9. All of the Windows operating systems are very hard-to-remove (you must wipe the hard drive),  behaving exactly like a virus or malware,  ever since Windows 95.  The last semi-stable version I had was Windows 3.1,  but it was nowhere near as stable as DOS 2 or DOS 6,  which never corrupted themselves or required reboots.

#10.  The Windows operating systems are all self-corrupting,  and they do not clean up the messes they make,  which clog up your hard drive memory,  and bog down your machine’s operating speed.  DOS did not do that.

Machine Guns in Las Vegas?

Update 10-3-17:  in red text below.

Update 10-4-17:  in blue text below.

Update 10-6-17:  in purple text below.

Under federal law,  a “machine gun” is a firearm that shoots more than one bullet per trigger pull.  The synonym for this is “fully-automatic”.  A “semi-automatic” weapon is one that sends one bullet per trigger pull,  loading the next round automatically.  If it doesn’t load the next round automatically,  that means the user must operate some sort of manual bolt or other mechanism to load the next round.  Bolt-action rifles,  pump or breakdown shotguns,  and ordinary revolver handguns fall into that last category. 

The M-16 used by US armed forces is indeed a machine gun,  a fully-automatic weapon,  although it can be operated as a semi-automatic single-shot weapon as well.  The same is true of the Russian-developed Kalashnikov AK-47.  These are true “assault weapons” for military use precisely because they really can be machine guns.  A military unit not so armed is at a lethally-distinct firepower disadvantage when confronted by such weapons. 

The AR-15 (and most modern hunting and sport guns) is a semi-automatic weapon,  not a machine gun / fully-automatic weapon.  The fact that an AR-15 looks exactly like an M-16,  has absolutely nothing to do with its rate of fire.   Calling it an “assault weapon” is wrong,  because no military unit today would ever go into combat with the AR-15.  They would be totally outgunned by any group with fully-automatic weapons.  It’s not about what the gun looks like,  it is entirely about what the gun can actually do.  Simple common sense.

Civilians in this country currently can indeed own or possess machine guns,  but what devices they can own,  and what they can do with them,  is very,  very,  very severely restricted.  This began with the National Firearms Act (NFA) of 1934.  That law came about because the mafia was causing mass death in the streets with the venerable old “Tommy gun”,  which really was a machine gun.  It severely restricted civilian ownership of fully automatic weapons,  short-barrel rifles and shotguns,  and certain explosives.  It was amended in 1968 and again in 1986.

The 1986 amendment restricted civilian ownership of fully automatic weapons to only those made before 1986,  only with payment of a $200 tax along with an enormous and very invasive application,  and only with a very,  very thorough ATF background investigation,  plus requirements for notification of the ATF any time the owner traveled with any of those devices. 

Such devices could not be updated or repaired with modern parts.  Parts for such devices are largely out-of-reach of all but the richest today.  There are no exceptions to allow for the ownership of anything newer than 1986.  There are no exceptions to any of the other requirements.

This status was superseded for a while in 1994 to disallow entirely the civilian ownership of those pre-1986 machine guns,  short-barrel guns,  and devices,  but that restriction expired in 2004.  So,  we are still under the 1986 version of the law today.

In all 50 states,  it may indeed be legal to own machine guns,  but only in accordance with the federal law!  If the possession or use is not in accord with federal law,  then such possession or use is presumed illegal under state law,  period!  Some states impose further restrictions,  some do not.  And that federal law is exactly the 1986 update of the 1934 NFA law.  Period.  No exceptions.

Modifying a semi-automatic weapon into a full-automatic weapon is indeed possible,  but it is generally not very easy to do.  It requires appropriate tools and knowledge and experience.  It also requires testing.  This is already illegal under any circumstances,  no exceptions. 

Update 10-4-17:  Two new technologies for increasing firing rate have come to light.  These are the "bump stock" and the "gat-crank".  These act to increase the firing rate of a semi-automatic weapon to that of a fully-automatic weapon,  without modifying the loading mechanism inside the weapon.  These are therefore technically legal,  but they definitely do violate the intent of the 1986 prohibition on all but grandfathered machine guns.  In my opinion,  this is cheating,  and should not be allowed.  

What the shooter in Las Vegas did,  and what motivated him,  are still the subjects of investigation.  Nothing is yet known with any certainty,  and such certainty is unlikely for quite a while yet. Update 10-6-17:  information in news reports keeps surfacing that point to mental illness of some kind in this shooter.  He got his guns legally,  because no judge ever had him committed.  If you look at the earlier article cited below,  that "leak" of guns into the hands of crazies is the most common cause of these mass shooting incidents!  

The best speculations are (1) he sneaked some 10 (weapon count has been climbing in subsequent reports,  both in the hotel and at his home) long-barrel weapons into his hotel room overlooking the outdoor concert venue,  (2) at least some of those weapons were machine guns based on the high rates of fire evident from the audio recordings of the event,  and (3) he fired into a dense crowd that could not move quickly,  so that without aiming,  he was certain to hit lots of people. 

Item 3 means that fully-automatic weapons are not required to exact a huge death toll,  but they do considerably raise it.  Not even semi-automatic weapons are needed.  A considerable death toll could still be expected with just single-shot,  bolt-action rifles.  So,  it’s not really about the gun,  it’s much more about the situation:  a densely-packed,  immobile crowd as the target from a nearby high place. 

Every time there is such a mass shooting event,  there is an immediate knee-jerk reaction:  a call for tighter gun control.  Always the same things are proposed,  and almost none of them would have prevented any of these events,  including this one!  The exceptions are (1) selling weapons too easily to crazy folks,  and (2) loopholes to the required background checks we already have. 

The problem here really isn’t so much the guns,  it is what motivates people to want to kill their neighbors.  What causes that?  I have never heard a good answer to that question.  Maybe it is past time to go find out. 

Update 10-3-17:  To find out what the gun violence is really trying to tell us,  go see my analysis of excerpts from the Mother Jones gun violence database.  It is not what you think!  This analysis is in the article titled "What the Gun Violence Data Really Say" dated 6-21-2016 on this website.  It has a list of titles and dates for other articles I have also written on this subject.  The navigation tool on the left gets you there most easily.  Click on the year,  then on the month,  then on the title. 

For those unwilling to go to the cited article and examine the data for themselves,  here is the short form of the message:  (1) we have a major "leak" of guns legally sold to people who are mentally ill,  but have never been so ruled by a court,  (2) we have a major problem with inadequately-defended (or entirely-undefended) gun-free zones,  which also invite terrorist attack,  and (3) the "usual" gun control proposals of "assault" weapons bans,  clip size limits,  and the like,  have already been tried and were already found to be ineffective.  

It's both that simple and that ugly.  Fix those two items properly,  and it looks to me like most of this problem goes away.  Item 3 tells you what not to do.  Update 10-4-17:  I also recommend outlawing "bump stocks" and "gat-cranks".  That won't prevent the incidents,  but it will reduce the death tolls.