Thursday, May 11, 2017

President Trump is Not Whom You Should Fear The Most

A lot of people are worried sick about what President Trump might do next.  Rest assured,  there are limits to the damage he can do.  But,  his term or terms in office will be quite interesting,  to say the least.  There is an old Chinese curse that is,  sadly,  applicable:  “may you live in interesting times”.

The office of President is specifically designed to limit the bad the President can do,  which also limits the good he or she can do.  All our Presidents do both good and bad while in office.  That’s just human nature.  The ones we consider "good" Presidents do more good than bad,  and vice versa.  It's never “either-or”.  

In Mr. Trump's case,  I fear he will do more bad than good,  not so much out of malicious intent,  but out of egregious ignorance compounded by unwarranted arrogance,  and a demonstrated inability to speak or value the truth.  But only time will tell. 

Age may not be the sole Constitutional qualification to be President that we need anymore.  Ignorance so egregious should be a disqualifier for the Presidency in the Constitution.  I’m not sure how you measure it to judge,  though.

The President Is Not the Only Problem We Face

Meanwhile,  I am more concerned about the damage the Republicans will do while in control of all 3 branches.  They are demonstrably unfit to govern,  since for more than 3 decades now,  the entire party has been held hostage to an extremist minority within their midst.  As the popular vote indicates,  the Republicans are about half of American voters.  So minority within their party is a small minority indeed,  compared to all of us.

That minority has managed to either purge,  or forcibly convert the rest of the Republican party into voting with them,  via the "primarying" threat.  Because it’s an extremist minority view that controls,  everything the party has ever claimed about economics and public policy has proven to be a lie,  for many years now. 

This is an example of a minority takeover of a government,  via extortion and propaganda.  This a minority uninterested in doing things for the people,  or even in maintaining their freedom,  despite what is claimed.  It is all about obtaining and wielding power over others.

Prior examples:  the Nazis,  the Bolsheviks,  the Chinese communists,  Pol Pot in Cambodia,  etc.  And THAT is why I am concerned about our future as a democracy.  Mr. Trump's deficiencies are just incidental to THAT threat.  

So Where Is the Alternative?

The problem is that we have little in the way of an alternative that is not just as dangerous.  The Democrats also sold out many years ago to the same financial giants as the Republicans sold out to,  more than a century ago.  They went "rich elite",  and completely forgot about their core constituency,  the lower middle class working people who typically held factory jobs. 

Which neglect is exactly why that base revolted and elected Mr. Trump!

The fundamental trouble is with party agendas on both sides:  they have little to do with reality,  being extremized WITHOUT REGARD TO FACTS,  just to differentiate the parties.  Political agendas do not make good public policy.  They never have. 

An Unexpected Problem Confronts Us

Money is power.  The love of money and power corrupts everything it touches,  as any preacher can tell you.  The real problem:  the financial giants own both parties,  and controlled who actually ran for President in 2016. 

Mr. Trump was the only Republican a “corrupt elite politician” like Hillary Clinton could conceivably beat.  And she did,  in the popular vote!  On the other hand,  Mrs. Clinton was the only Democrat that an ignorant incompetent like Mr. Trump could possibly beat.  And he did,  in the electoral college!  PER THE PLAN!

Both party primaries were deliberately rigged for these two candidates to keep the "blood and circuses" interesting enough to prevent public perception of who was really pulling all the strings behind the scenes.  You didn't think Mr. Comey was incompetent enough to violate FBI investigation procedures by making those two public announcements (July and October) about Mrs. Clinton's email-mishandling,  did you?  And it WAS mishandling,  but she is not unique in that fault,  not by a long shot. 

Mr. Comey was forced to do those irregular election-swaying announcements by the financial giants to keep the race looking "even" for a longer time,  until the giants could get their man in at the last moment.  This was with Russian fake news (via Wikileaks,  Sputnik,  and RT) helping to sway voters' minds,  in addition to Mr. Comey’s October announcement. 

The collusion deserving investigation is less about Mr. Trump's campaign than it is those financial giants.  They wanted an incompetent they could influence or control to help give them everything they want.  They got him. 

Can We Fix This Problem?

You cannot expect a wholly-bought Congress to change this,  no matter which party is in power.  They work for who bought their jobs for them,  not us.  This has been true for decades now.  They would have to cooperate with a grassroots Constitutional amendment process,  too,  and they will not,  because of the money. 

Ballot-box solutions are thus pretty-well ruled out.  This evil takeover of our government will not go away without some sort of revolution,  I fear.  The giant corporations and banks control the entire federal government,  which means they control the armed forces.  And they know the rebellion is inevitably coming. 

Why else would EPA enforcement officers (and similar non-military career government workers) need body armor,  swat vehicles,  machine guns,  and most of the ammunition being produced?  They are to serve as adjunct troops to help put down the rebellion!

Compared to this picture,  an arrogant and ignorant incompetent like Mr. Trump is small change,  even with his finger on the nuclear trigger.  

Keep your guns and ammunition hidden.  You’re going to need them.  Just as the Second Amendment says. 


Saturday, April 15, 2017

Do We Fight Global Warming Or Not?

Note:  article was updated 4-23-17 in purple text below to include sources of traceable data.  

Note:  another update added below in blue text 4-25-17.

Note:  one slight edit adds an item in red text below 5-4-17.

This is an issue that has become politicized to the extreme,  which precludes rational action. What I present here has absolutely zero to do with ideologies or politics.  It is simple logic and common sense. 

There are two things to consider,  but only one available choice.  Whether humans cause global warming or not is not a matter of choice,  it is something decreed by nature,  which does not tell us which is true.  Our only choice is whether or not to act,  based on what we do know. 

What we know is this:  (1) there is a huge volume of ice on Earth located above sea level,  (2) if even some of it were to melt,  sea levels would rise sharply,  (3) added heat melts ice,  and (4) most of our critical institutions and a major fraction of our population live in the zone threatened with flooding. 

What portends here is a disaster far exceeding the temporary flooding of a city by a hurricane,  or the migration of millions out of Syria and Africa to escape war.  What could happen is the forced migration of billions,  and (nuclear) war over failing food resources.  So,  this decision is important to get “right”. 

Filling Out The Decision Matrix

One simple way to decide this is by a version of the trade study matrix,  a pretty standard tool.  However many choices you have is the number of columns (in this example 3),  however many versions of the unknown natural issue there might be is the number of rows (in this example 2).  That gives you a 6-hole pigeon-hole matrix to fill in with likely consequences. 

There are two rows because human emissions might,  or might not,  cause global warming.  You do not get to choose between them;  this is decided by nature,  not humans. 

There are three columns instead of two,  because if we decide to act,  there’s two ways this action might turn out.  There is only one,  if we choose not to act.  Acting versus not acting is the choice available to us.  If we act and it doesn’t work,  we’d better already be working on how to cope (the third column).

As for the consequences,  they need not be detailed,  and it is OK to exaggerate them for better contrast.  

If we choose to act,  we will spend lots of money to act,  and there will be monetary losses,  too.   These costs could range from significant (damaged economies) to catastrophic (going back to the stone age).  That variation doesn’t matter,  just fill in all four “choose to act” cells with “lose $$”.

If we choose not to act,  then the consequences depend upon what nature does not tell us:  whether or not human-caused global warming is real.  If not real,  there will be no meltdown,  no sea level rise,  no migrations,  no war,  and no money lost.  If real,  all those things will happen,  and both money and lives will be lost (at catastrophic levels). 

That fills in all 6 cells with consequences.  5 of the 6 involve lost money,  there is no avoiding that.  1 of the 6 involves life loss as well as loss of money;  that one is really bad.  1 of the 6 has no bad consequences in it at all. 

Now We Must Choose

You cannot choose which row you want (political ideologies notwithstanding).  You can only choose a column!  The standard way to use the matrix is to pick the outcome that you cannot abide,  and then cross out the entire column that contains it. 

In this example,  losing lives is to be avoided,  which rules out choosing not to act.  This valuing of lives over money is in accordance with the teachings of all 3 Abrahamic religions in the West.  Most of the Eastern religious traditions agree. 

That result says:  act,  and be prepared ahead of time to cope,  if your initial action fails. 

Did you notice that not once did I refer to any of the prognostications or temperature history data of the climate science community?  I didn’t need it to make this decision.  I need it only to help define the actions we might take to mitigate this threat:  reduce greenhouse gas emissions. 

And,  there is another independent science dataset that says the same thing:  observed ice melting behavior as the fossil fuel-guzzling population has exploded.  Getting the same answer by two independent means lends a lot of confidence to that answer. 

Update 4-23-17:  Sources of Real Data to Consider

There are ice core data that cover atmospheric composition during the ice ages and the warm periods in between.  This is based on the actual composition of the ancient air trapped in the bubbles in the ice.  The atmosphere is mixed well enough that this composition is not restricted to polar regions,  it is global.  These can be dated by the layers,  similar to tree ring dating.  Here is that data for atmospheric carbon dioxide over the last 400,000 years,  obtained right off a NASA website:

You can see the 4 dips to 180-200 ppm at the height of each of the 4 main glaciations of the ice age. We know when these glaciations occurred from the timing of the evidence in the rocks:  they show marks of glacier passage,  and the debris left behind on melting.   Note that it never got above about 280-290 ppm during the deglaciated warm intervals.  Ancient is to the left,  modern is to the right.  You can even see the little "wiggle" in the curve at about 260 ppm about 10,000 years ago that is the sudden cool-down they call the "Younger Dryas".  

From 180 ppm to 290 ppm encompasses atmospheric composition all the way between fully glaciated to fully warm.  Correlation does not establish causality,  that has to come from elsewhere (such as basic demonstrable physics).  So,  is something else going on?  Such as Milankovitch orbital cycles?

The thing we have that best models the cycling of the ice ages is Milankovitch orbital cycles.  This is not really a fully causal model,  except for the notion that more sunlight striking northern hemisphere land leads to warmer conditions globally.  It pretty much correlates with the advance and retreat of the ice;  not perfect,  but very,  very good.  It is limited;  for one thing,  the distribution of continents was different millions of years ago.  

The basic physics is simple:  ice melts if heated.  The Earth's "average" temperature is an energy balance between lots of visible and ultraviolet light coming from the sun,  and some heat of radioactive decay and original formation escaping from the interior,  versus the infrared heat re-radiated back out into space by the warmth of the Earth's surface.  And,  extra carbon dioxide in the atmosphere interferes with that re-radiation,  because it is less transparent to infrared than oxygen and nitrogen,  so the surface must warm further to radiate against the resistance of the carbon dioxide.

You can verify this effect for yourself without actually doing sophisticated measurements:  simply set two bell jars covering thermometers out in the sun at the same time.  One has air,  the other you fill with carbon dioxide (the extreme case).  Both thermometers rise.  But,  the carbon dioxide-filled jar's thermometer will read a lot higher than the air-filled jar's thermometer.  Both are "greenhouses",  but the carbon dioxide gas is far more potent as such than oxygen and nitrogen in the air.  

Another version of that very same chart I obtained from Wikipedia,  to which an inset was added showing atmospheric composition over only the last few centuries.  This makes the point that our unburying of carbon-containing fossil fuels and releasing it as exhaust gas carbon dioxide,  has had effects since the start of the Industrial Revolution,  and really sharp effects in the last 5 decades as our population explodes exponentially.  

If you look on much longer time scales,  there are other things going on as well.  On a time scale of 100 million years,  the astrophysicists tell us the sun has brightened by 4% or thereabouts.  On a 4.6 billion year time scale,  they tell us it has brightened by about 30%.  

Before about 380 million years ago,  there was no life on land.  Before 600 million years ago,  there was only single cell life in the ocean.  Before about 2.5 billion years ago,  there was no oxygen in the atmosphere.  And who knows what the surface air pressure was during those times (which also affects how good a "greenhouse" it makes)?

All we know is that there more carbon dioxide half a billion years ago than in "recent" times (only the couple of million years).  The sun was dimmer,  and yet the geology indicates ice-free conditions.  This chart was published a few years ago in the refereed journal "Science",  published by the American Association for the Advancement of Science.  It's based on atmospheric composition inferred from rock chemistry,  and it's pretty good back to the Cambrian,  570 million years ago.  Much before that,  it's inherently rather speculative,  which explains the scale change representing time.  These are indirect measures,  which explains the lack of scale tick marks on carbon dioxide concentrations,  which were roughly around 1000 to 2000 ppm during the Mesozoic.  

Update 4-25-17:

What the long-term carbon dioxide and temperature chart makes clear is twofold.  (1) Carbon dioxide fluctuations do not cause ice ages,  because there was little change in level during the Pleistocene Ice Ages,  and carbon dioxide levels were much higher during the earlier ice ages.  (2) Carbon dioxide in the air does indeed warm the planet,  as evidenced by ice-free intervals at high carbon dioxide earlier in Earth's history,  when the sun was significantly dimmer.  

Something else causes ice ages.  Many things,  this is poorly understood.  There have been many of these ice age events over geologic time:  the Pleistocene event we are most familiar with,  the Jurassic-Cretaceous event,  an event between the Carboniferous and the Permian,  another between the Ordovician and the Silurian,  and who-knows-what during Pre-Cambrian times.  

Climate-Modeling Science

The overwhelming majority of climate scientists agree that humans are causing major effects with greenhouse gas emissions.  They arrive at this conclusion with a combination of (1) computer modeling of climate,  and (2) various proxies for past temperature data earlier than historical measurements. 

There is inherently a lot of uncertainty in the computer modeling,  and a lot of inference in the proxies for past temperatures (unlike the ice core data for atmospheric composition).  There is potential for error,  disagreement,  and even fraud.  Many folks outside the community are uncomfortable with that,  and this is the weakness exploited by those who prefer to disbelieve that we are causing climate change. 

Ice Melting Behavior

Ice behavior is unambiguous.  The mountain glaciers have been generally receding since the 19th century.  Now there are enormous summer sea ice losses,  and thousands of summer meltwater lakes on Greenland,  that we have never seen before!  The co-timing of these symptoms with the increases in measured atmospheric carbon dioxide to unprecedented levels since 1958,  is quite damning.

There is a documentary film available in whole or in part on Youtube named “Chasing Ice”.  It was made by James Balog as part of his Extreme Ice Survey (EIS).  The award-winning film was first shown in 2013.  The time lapse photography of many glaciers' melt-back in the last 30 minutes of that film makes my point better than any words. 

Trade Study-Recommended Actions

The mitigation action to take first is to cut back carbon dioxide and methane emissions as fast as we can,  but without hurting or killing somebody for lack of energy, which limits how fast we can do this.  The coping action to take in case mitigation fails is twofold:  (1) start stockpiling foodstuffs,  and (2) to start moving critical institutions and assets to much higher ground.  

Any other “geo-engineering” activities we contemplate must be reversible,  because we simply do not know that they will do more good than harm.  If they do not work,  we have to be able to undo them. 

It is that simple. And it is that stark.  And,  it has absolutely nothing to do with politics or ideology.  Those who claim otherwise are lying to you.  Follow the money to see who and why. 

Previous Related Article on this Topic

There was one earlier article that I wrote on this topic,  which the current article updates and replaces.  That was “On Global Warming”,  dated 1-12-2010,  and sharing the same search keywords you can use to filter searches for this topic on this site:  "bad government",  “bad manners”,  “climate change”,  and “idiocy in politics”.  That older article was last updated in 2014 to show a simpler 4-cell version of the 6-cell trade study matrix presented here.   It now refers the reader to this article.  

Saturday, April 8, 2017

The Time Has Come to Deal With Iran and North Korea

Both of these rogue nations are pursuing ballistic missiles tipped with atomic weapons.  They have made enough progress that we should be seriously concerned,  especially in the case of North Korea.  Action is required now with North Korea,  and very,  very soon with Iran.

North Korean Progress

North Korea has made enough progress toward atomic weapons that they have been testing such weapons underground for several recent years.  They have been doing this for enough years to have at least begun (and possibly completed) the necessary miniaturization of the atomic weapons,  so as to fit a more ordinary-sized rocket.  It is the rocket that is still giving them problems,  so that many rocket flight tests have been made recently. 

There are 4 things the North Koreans require,  in order to strike a mainland US city with a blast weapon:  (1) a miniaturized atomic bomb,  (2) a reliable launch rocket,  (3) a heat shield for the warhead to survive re-entry,  and (4) guidance precise enough to actually hit fairly close to the intended target (both detonation altitude and miss distance are important). 

There are only two of these needed to damage us severely with the electromagnetic pulse (EMP) of a nuclear explosion in near space.  To do that requires only the bomb and the rocket;  the precision guidance and a heat shield are unnecessary. 

If they have not successfully miniaturized their atomic bomb yet,  they will within a year or two at most.  By then,  their rocket should also be flying reliably.  That means we are credibly at risk “right now”,  and very most certainly within a year or two.   It is now past time to put an end to their efforts.

Iranian Progress

Iran already has the rocket “in hand”:  they have launched satellites into orbit.  Our own history shows that any satellite launcher can fly sub-orbitally with a larger payload.  That payload can easily be a miniaturized atomic weapon. 

That is why Iran’s main effort in recent years was toward those atomic weapons.  Like North Korea,  they could do us great damage with only the rocket and bomb as an EMP attack.  With a heat shield and precision guidance,  they could also do a blast weapon attack.  Those last two components are easier to do,  than the miniaturized bomb and the rocket,  and easier to conceal. 

The nuclear “deal” with Iran has temporarily slowed (perhaps even halted,  but I really doubt that) Iran’s uranium fuel program.  We have already seen them highly-enrich uranium,  something unnecessary to run a power reactor.  Highly-enriched uranium is only bomb material,  simple as that. 

However,  not often considered in the news reports is the fact that even low-enriched uranium can be used in a modified reactor design that breeds plutonium from the non-fissionable leftovers from the uranium enrichment process.  Plutonium makes even better bomb material,  although how you set it off is different from highly-enriched uranium.  But,  how you set off either is publicly-available knowledge! 

The upshot of that is that any country who can build reactors that use low-enriched uranium,  can also breed plutonium and make plutonium-based atomic bombs!  The Iran nuclear deal does NOT prevent that from happening!  From that point,  all that is required is miniaturization of the atomic bomb to fit the rocket.  And Iran already has the rocket!

It might take Iran a single-handful of years to build plutonium bombs and get them miniaturized successfully.  At that point,  they can successfully strike us with atomic bombs.  It is therefore pretty-much time to put an end to their efforts,  too. 

What Could We Do About North Korea?

North Korea has a weakness we can exploit as a unilateral action:  their rocket is still unready to fly their atomic bombs.   Stop the rocket tests,  and you can still stop their capability to hurt us,  at least for a while.  Longer-term,  there must be regime change in North Korea,  or else this threat will never go away.   

We have various battlefield and longer-range anti-missile and anti-satellite weapons.  Some of these seem to work,  at least under restricted circumstances.  With all of them,  there is still credible doubt about their efficacy during general warfare.   But what we need here is only efficacy in a restricted circumstance:  shooting down every test rocket launch conducted by North Korea,  for the forseeable future. 

That is exactly what I propose as the initial step against North Korea:  shoot down every single test missile they launch.  This has two effects:  (1) North Korea cannot verify their rocket to be reliable,  at least for the short term,  and (2) it shows China we are very,  very serious about taking unilateral action if they do not rein in their protégé state. 

In the longer term,  we will need the help of China to resolve this situation.  They are the source of imports and support that actually keeps the rogue state of North Korea alive and functional.  It is in China’s interest as well as ours that there not be a failed state in North Korea.  Further,  there is some reason to believe that the Kim dynasty in North Korea has limited days left.  When it ends,  chaos is the most likely result,  unless a major power steps in. 

But,  I rather doubt that China might support reunification of the Koreas under the government of South Korea,  even though that would be a favorable outcome for them and for us.   So,  the realistic prospect is that there will still be two Koreas indefinitely into the future.  The “trick” is getting China itself to replace the paranoid Kim dynasty with something more sane and more tolerable,  to us and to them.   

What Could We Do About Iran?

This is by far the tougher problem to solve. 

Iran has the rocket,  but they do not yet have the bomb to ride that rocket.  It is only a matter of a very few years before they do have the miniaturized bomb,  despite the nuclear deal.   Whether they cheat on the deal,  or not,  makes no real difference. 

The exact locations of all their nuclear facilities are too uncertain for us to strike,  and those we do know precisely,  are buried deep underground.  Conventional weapons simply cannot take them out;  only a ground-penetrating nuclear strike could do this job.  The world will not condone that. 

Like North Korea,  Iran is ruled by extremists who policy objectives are demonstrably insane by any standards that we in the west understand.  In that respect,  they differ in no practical way from ISIS,  Al Qaeda,  or the Taliban. 

Unlike North Korea,  Iran has no major power as a “sponsor” to keep them functional.  In point of fact,  Iran is a major regional power all on its own,  complete with proxy armies (Hezbollah,  Hamas,  and some others) to do its bidding to spread chaos everywhere. 

Diplomacy (the nuclear deal) has slowed the problem only a little,  but definitely has not stopped it.  Short of nuclear genocide,  there is little we the US can unilaterally do,  or even do with multiple allies.  Yet something must be done,  and all the Iranians’ neighbors agree.  The people of Iran are actually good and decent folk;  they do not deserve nuclear extinction.  But their government certainly does!

This one is a real “rock-and-a-hard-place” problem.  About the only hope I can offer is that diplomacy with Iran might be more effective,  if we have already made an example of North Korea.  And also perhaps of their co-supported (with Russia) puppet:  Assad in Syria. 

To that end:  put an end to Kim Jong Un in North Korea,  then make sure Bashar Assad in Syria dies for conducting chemical warfare attacks.  Target him (instead of airfields) with cruise missiles.  Let the Russians install whomever they want in Syria to replace him,  but we must be sure Assad dies.  Period. 

That is a very difficult prescription indeed,  but it must be done!  After it is done,  both us and the Russians may actually benefit.  And those extremists ruling Iran may be more tractable. 

Maybe.  Maybe not.  No guarantees. 


Previous Related Articles That This Article Supersedes:

date,  search keywords
article title

4-6-09,  North Korean Rocket test             
Thoughts on the North Korean Rocket Test And Beyond      

12-13-12,  current events, North Korean rocket test                   
On the 12-12-12 North Korean Satellite Launch

2-15-13,  Mideast threats, North Korean rocket test
Third North Korean Nuclear Test
4-5-13,  current events, North Korean rocket test                  
North Korean Threat Overblown, So Far

9-12-15,  bad government, bad manners, current events, idiocy in politics, Mideast threats                
Iran Nuclear Deal Nonsense


Sunday, April 2, 2017

Spacex Re-Flies Used Booster

Update 5-4-17:  They've done it again,  this time launching a recon satellite for the military,  and once again they successfully recovered the first stage booster for reuse.  This was their first launch in the launch business sector that was previously a de-facto monopoly for ULA.  

Original Article:

Spacex’s seemingly-routine successful launch on 30 March 2017,  of a satellite to geosynchronous transfer orbit is a bigger deal than it first seems.  The Falcon-9 first stage booster rocket was a used item landed before,  and now landed again.  

This has never-before been done with a rocket capable of reaching orbit.  It portends a near-term dramatic drop in launch costs to space,  but only if the technology proves out the way hoped. 

Reusable launch to orbit is more demanding than reusable launch into suborbital flight.  This is because the conditions when the used stage returns to the atmosphere are far harsher for an orbital launcher. 

In the suborbital arena,  it isn’t widely publicized yet,  but Blue Origin has flown and landed the same New Shepard booster rocket some 5 times now.  That,  too,  has never before been done.  

More To Come

There’s more reusable firsts waiting in the wings.  Virgin Galactic’s “Spaceship Two”,  and XCOR Aerospace”s “Lynx”.  Both are reusable suborbital spaceplanes for the tourist industry. 

There is also Sierra Nevada’s orbital “Dreamchaser” spaceplane,  which is undergoing its initial tests.   And,  Spacex’s “Dragon” capsules were designed from the beginning to be re-used. 

Blue Origin does plan to enter the orbital market with its New Glenn rocket,  which is to be reusable.  That activity is just beginning:  the new big engine for it has begun ground testing. 

Spacex plans to fly soon a much larger reusable rocket called Falcon-Heavy,  that is based on its Falcon-9 hardware.  As I understand it,  the first flight of Falcon Heavy is deferred,  until pad 40 at Cape Canaveral is repaired after last year’s explosion.  This is wise,  because a problem test-flying Falcon-Heavy off of pad 39 then cannot stop commercial Falcon-9 launches off pad 40. 

Spaceship Two is about to resume flight testing after an accident in test a while back.  It will carry about half a dozen passengers,  and a crew of two pilots.  It is launched from a large carrier plane. 

Lynx is smaller:  one passenger,  one pilot.  But it simply takes off from a runway,  flies into space,  and returns to that runway,  perhaps up to 4 times per day.  The first flight test article is nearing completion. 

A Look Behind the Curtain

What is so remarkable about all this is not so much what is being done,  but who is doing it!  Not one of those names is part of the industry that worked with NASA or the USAF all these decades,  which is sometimes called “old space”.  This is entirely what we might call “new space”. 

That is not to say that these “new space” companies don’t work with NASA or USAF,  because many do.  Spacex has NASA contracts to deliver cargo with its Dragon to the space station,  as does Orbital ATK with its Cygnus. 

Spacex has a NASA contract to develop a crewed variant of “Dragon” to deliver astronauts to orbit,  as does Boeing with its CST-100 “Starliner” capsule.  Sierra Nevada’s “Dreamchaser” spaceplane was also contracted by NASA for this,  got dropped for a while,  but may now get contracted again. 

NASA and USAF are also interested in Blue Origin’s New Glenn orbital rocket,  especially the engines that will push it.  These engines may be candidates for a follow-on launch rocket to the Atlas-V,  that both agencies routinely buy from the Boeing / Lockheed-Martin venture ULA (United Launch Alliance). 

And Spacex now has contracts from USAF to launch some of its satellites.  Before,  only ULA had any of that business. (See update 5-4-17 above.)

Why This Is Happening

What broke this market open for “new space” entrants was their lower prices.  And that came from competition in the commercial satellite launch business,  something neither NASA nor USAF does.  These American companies and a variety of foreign companies all had to learn how to reduce price,  competing in that growing commercial satellite business. 

Up to now,  the “secret” was simplification of the logistical “tail” that supports production and flight of these rockets as expendable vehicles.  Reducing that support tail from the size of a major city to the size of a small town reduced the “typical” per-launch price from many hundreds of millions of dollars to only several tens of millions of dollars:  almost a factor of 10.

Reusability promises to reduce that by at least another factor of 10!  Maybe more!

It is fundamentally the large size of the commercial satellite launch market that can support so many companies competing in it.  That has taken decades to grow.  But the competition that is inherent with many companies is what spurs the innovation that cuts costs,  allowing lower prices. 

Consolidation into one,  or a few,  stops that competition.  That kills downward pressure on prices,  and thus dis-incentivizes innovation.

Amazing what the truly-competitive market can do for you,  given time and opportunity.  

Thursday, March 23, 2017

Water on Mars?

This is only meant as a funny.  The caption is something like "first discovery of water on Mars".

Saturday, March 18, 2017

Bounding Analysis for Lunar Lander Designs

I did this as a "clean sheet" bounding analysis.  Friends I correspond with have asked repeatedly how a lunar base might be established,  and with what.  I know the most about Spacex rockets and capsules,  but actually fleshing out these designs could use anybody's existing equipment,  in whole or in part.  The challenge I now throw out to them is to design something within these limits.

Bounding Analysis for Lunar Lander Designs  
GWJ  3-18-17 completed 3-18-17

The scenario here is a lander delivered “neat” to lunar orbit as an unmanned item.  A crew will arrive separately to rendezvous with it in lunar orbit.  The plane of that orbit is presumed to be very close to the ecliptic.  Orbital direction is retrograde,  in accordance with the figure-eight patched-conic trajectory used during Apollo.  The delta-vee to land one-way is 1.68 km/s.  For design purposes,  a few percent higher is used to provide a little margin:  1.75 km/s. 

The lander is delivered “neat” to lunar orbit,  meaning the rocket that takes it to the moon must do the “burn” to put it into lunar orbit.  The total rocket design delta-vee from the surface of the Earth to do that is at most 12.4 km/s,  when the first 8 km/s to Earth orbit is factored for drag and gravity losses by 1.05.  This is very close to the surface launch for a more-or-less worst-case slow trajectory to Mars,  which is about 12.1 km/s,  factored the same way.  That way,  the tonnage sendable onto a Mars transfer trajectory is almost the same as what can be delivered into lunar orbit,  for our purposes here. 

Descent Design Requirements

Spacex lists on its website that its Falcon-Heavy can send 13.6 metric tons to Mars,  flown fully-expendably,  for about $90 M launch price.  This heavy-lift booster hasn’t yet flown,  but it should fly this year (2017).  Reducing that payload slightly for the slightly-higher delta-vee to lunar orbit,  call that a max payload to lunar orbit of an even 13 metric tons.  

For the descent stage,  ready to fire in lunar orbit,  we are looking at an ignition mass of 13,000 kg maximum,  and a required design delta-vee of 1.75 km/s.  Propellants should be storable,  since days to weeks,  even months,  in space (or on the moon) are contemplated.  With nozzles designed for vacuum,  and assuming NTO-MMH propellants,  a delivered Isp = 335 sec is quite realistic.  Engine thrust/weight ratio of 100 Newtons-of-thrust per Newton of engine Earth weight seems feasible.  

Thrust to ignition Earth weight ratio should just barely exceed lunar gravity’s pull,  so that plenty of thrust margin is available at burnout weight:  0.2 seems “reasonable”.  We’d like the vehicle acceleration at burnout to be less than 1 gee,  preferably under 0.5 gee,  to keep the ride from being too rough,  and to limit throttleability requirements to feasible values.   

The propellant tanks will need a sun-reflective surface and some insulation,  plus electric in-tank heaters,  on a single-hull tank.  That means the descent stage propellant tankage will not be quite as lightweight as that of an expendable booster.  Just considering the tankage alone,  a 95-5 split of propellant to tank masses seems reasonable to assume (Wp/Wt = 95/5 = 19). 

The rest of the stage structure must bear the thrusted flight maneuvering loads carrying as large a payload as possible,  plus incorporate a set of broad-span landing legs,  and some means of unloading large items (ramps,  crane,  etc.).  An inert structural fraction for the stage near 15% should cover all of this.  That fraction does not include tank inerts or engine hardware.  Those get figured separately,  and then added to determine an overall stage inert mass fraction. 

The objective here is to determine max payload mass within that ignition mass limitation.  That payload can be either (1) cargo delivered one-way,  or (2) an ascent vehicle carrying minimum crew and cargo weight.  They mass the same,  though. 

Sizing a “Clean-Sheet” Bound on the Descent Stage

Exhaust velocity is rather accurately estimated as Vex, km/s = 9.8067*(Isp, sec)/1000.  That and the design delta-vee value combine to determine stage mass ratio MR = exp(dV/Vex).  The required propellant fraction (of ignition mass) is Wp/Wig = 1 – 1/MR.  The corresponding fraction for tankage inerts is Wt/Wig = (Wp/Wig)/(Wp/Wt).  The corresponding engine inert fraction of ignition mass is We/Wig = (T/Wig)/(T/We).  The rest of the stage structural and equipment inerts is represented by the 15% figure.  These total together for the overall stage inert fraction. 

Payload fraction of ignition mass is just 1 minus the propellant fraction and minus the sum total inert fraction.  Once you specify an absolute payload mass,  that determines ignition mass,  from which all the component masses are determined by their fractions.  That finalizes the weight statement.  For this bounding exercise based on Falcon-Heavy delivery,  those results are in Figure 1.  

Figure 1 – Limits for Descent Stage,  One-Way,  Falcon-Heavy Delivery to Lunar Orbit

Payload is 5.372 metric tons.  This could be all cargo,  or it could represent a crewed ascent stage.  If cargo,  that’s $90M/5.372 metric tons = $16.8M per metric ton delivered to the surface of the moon.  Actually,  you design to a slightly-smaller payload mass,  because of all the uncertainties.  There is always the unexpected outcome,  when sizing vehicles like this “from scratch”.  The weight margins don’t have to be all that large,  because I already put that into the design delta-vee figures. 

Ascent Design Requirements

The same propellant and tankage choices are presumed.  The same engine T/We is assumed.  A slightly-higher T/Wig = 0.3 is assumed,  to accelerate “smartly upward” against lunar gravity.  Stage inert fractions can be lower since no unload equipment or landing legs are needed.  However,  these inerts are likely higher than a typical booster rocket (5%) because of the protective cabin surrounding the crew,  the docking hatch,  and the instruments and controls they must use.  I simply assumed 10%. 

This ascent stage must ascend to lunar orbit (requiring 1.68 km/s),  and also maneuver to rendezvous with the crew return craft left in lunar orbit.  It therefore needs more design delta-vee than the descent stage.  Call it 2.0 km/s,  for a kitty of 0.3 km/s to cover maneuvering and the unexpected. 
Its maximum ignition mass cannot exceed the descent stage payload capability of 5372 kg.  Prudence dictates very slightly less.  Call it 5360 kg for design-bounding purposes. 

Sizing a Clean-Sheet Ascent Design to Fit the Descent Stage

All the calculations and equations are basically the same as before.  I simply used the same spreadsheet with different numbers.  The results are given in Figure 2.  Deliverable “payload” is 2235 kg,  which would be suited crew plus a few of days of life support,  plus any samples sent back to Earth.

Figure 2 – Limits for Ascent Stage,  One-Way,  To Fit Descent Stage That Falcon-Heavy Can Deliver

For the sake of argument,  use 80 kg per person body weight,  and 120 kg for a surface EVA-capable pressure suit.  That’s 200 kg per person.  Set food,  water,  and breathing oxygen supplies to 100 kg to cover an unexpectedly-long rendezvous interval of several days.  That’s 300 kg allotted per person.  There’s “room” for 7 such masses in the payload. 

If this were 6 crew,  there’s room for around 300 kg of samples or return cargo.  If the crew is 5,  there’s room for about 600 kg of samples or return cargo,  and so forth.  But the point is,  there’s room for a much larger crew than Apollo had.  That’s partly the difference in technologically-achievable storable propellant performance,  and in structural technologies,  since the 1960’s.  The rest is landing without unknown obstacles in your path,  which is what happened on Apollo 11,  nearly depleting its propellant. 

How This Can Be Used

The one-way cargo-only variant can be used at $90M a shot to deliver 5.36 metric tons of cargo to the moon ($16.8 M/delivered metric ton).  Several could be sent to the same site.  Some of these could be the modules from which some sort of surface habitat could be assembled.  The rest could be the supplies,  equipment,  and surface rover vehicles needed to operate that base. 

The manned lander conforms to the same 5.36 metric ton weight limit.  If crew were 3,  then 1200 kg of surface supplies could go down with them.  If crew were 2,  then 1500 kg of cargo could ride down.  Reducing the ascent load just increases the rendezvous maneuver capability upon returning to lunar orbit,  a very beneficial safety factor.    

Say,  we sent 9 of these to the moon:  6 cargo-only landers and 3 landers with manned ascent stages,  each with a crew of 2 and 1500 kg of cargo on board.  That gives us three ascent vehicles on the lunar surface ready to use,  when the entire crew really only needs one to return.  Added safety,  that is. 
That’s a total of 32.16 tons delivered with the cargo landers,  and 4.5 more tons sent down with the manned landers,  for a total crew of 6.  Assume simply for the sake of argument that the surface habitat requires 20 tons.  We need to reserve 0.6 tons of supplies for the crew to ascend.  Assume two rovers,  each 1 ton.  Assume one electric backhoe-like device,  at 2 tons. 

36.66 tons total delivered cargo,  less 20 ton habitat,  4 tons for vehicles,  and 0.6 tons for ascent supplies,  leaves 12.06 tons allocatable for surface stay supplies and other equipment items.  At a nominally-assumed 10 kg life support per person per day for 3 months,  then about half that 12 tons is something other than life support supplies.  Also nominally,  3 months of life support supplies for a crew of 6 is pretty close to a lander’s deliverable payload at 5400 kg.  I tried to overestimate this requirement. 

Looks to me like there is very good potential for establishing a fairly substantial lunar experiment station,  temporarily occupied for a considerable time (at least 3 months).  This requires 9 Falcon-Heavy fully-expendable launches for the landers,  plus one more to send the crew in a crew Dragon (with its trunk modified to carry propellant,  something not addressed here),  for $900 M in launch costs.  If launch costs were 20% of the program that develops these vehicles and the surface equipment,  total program cost to put a small base temporarily on the moon would be in the ballpark of $4.5 B. 

Launching another cargo lander every 3 months or thereabouts brings the supplies to keep that base permanently occupied at crew size 6.  Maybe switch out crews yearly,  by adding a crewed Dragon to lunar orbit along with a fresh manned lander to take them down to the surface.  That’s a total of 6 Falcon-Heavy launches per year to maintain a continuous presence at the base.  That’s $540M per year to maintain the base after it is built,  plus the costs of keeping the necessary vehicles and equipment in production.  Development is complete,  so call launch costs ~50% of continuing program costs. 

About $4.5 B to establish a 3-month-capable,  6-man base on the moon,  and about $1B/year to keep it continuously manned and operating is just not very expensive as space ventures go!  This analysis is based on the use of a commercial heavy lift rocket that is far less expensive to use than NASA’s SLS,  and which will also be far more available for routine use multiple times per year,  than NASA’s SLS ever can. 

Blue Origin is also planning to get into this kind of lunar capability with its New Glenn rocket.  Between them and Spacex,  putting a base on the moon looks to be quite feasible and quite affordable.  This could provide the bootstrap start needed to begin doing something useful,  or for profit,  on the moon. 

Final Remarks

This kind of experiment station allows evaluation of low-gravity effects upon health versus the zero-gravity effects that we are familiar with in Earth orbit.  It allows a place to experiment with increasingly-capable recycling life support systems.  It allows a place to experiment with meteoroid and radiation protection by regolith cover.  It allows a place to experiment with ways and means to overcome contamination and wear issues with very-fine-but-sharp-edged dust particles.  All these are needed to visit Mars or the asteroids,  and are available on the moon “close by” in case of trouble. 

The same base allows experimentation with ways and means to dig and drill deep in a harsh environment.  It allows experimentation with the recovery of mineral resources.  It allows experimentation with how to establish roads under such conditions,  so that future long-distance surface transport becomes feasible.  These things are needed for establishing useful and prosperous industrial applications on the moon and Mars,  and to some extent the asteroids. 

This is the kind of thing we should have attempted to close-out Apollo,  had a useful lunar presence been the goal,  instead of “flags-and-footprints”.  It is still a good rationale for returning and doing something very much like what I described here,  as a first step. 

Addendum:  Crew Dragon Modified to Leave Lunar Orbit

The “design” trajectory to reach lunar orbit is pretty much the same as was used for Apollo decades ago.  A direct launch from Canaveral into low Earth orbit more or less eastward at low inclination (the part requiring factoring ideal delta-vee for gravity and drag losses),  followed by a burn to escape onto the lunar transfer trajectory,  and a final upper-stage burn to place the payload into a retrograde orbit about the moon.  The worst-case total rocket design delta-vee for this is just about 12.4 km/s (factored),  and worst-case 0.8 km/s to leave lunar orbit onto a trajectory home.  See Figure 3. 

If we stay under 13 thrown metric tons,  the Falcon-Heavy should have enough delta-vee capability to put that 13 tons into lunar orbit,  same as the lander designs just bounded above.  The problem is then leaving lunar orbit with enough propellant reserve to cover attitude control and a powered landing on land back on Earth (Spacex’s preferred mode).  Attitude control consumption should be modest,  but we might need around 0.5 km/s capability to land safely on Earth,  where capsule terminal fall velocities are only around half a Mach number.  0.8 + 0.5 + small change is close to 1.35-1.4 km/s delta vee capability demanded of the Super Dracos on crew Dragon.  It simply does not have that much capability without extra propellant added in the trunk,  and connected to the system in the capsule.

Figure 3 – Design Trajectory and Delta-Vee Requirements

Design Requirements for Modified Crew Dragon

Total delta-vee capability 1.35 km/s min,  1.40 preferred.  Maximum spacecraft mass at launch 13.0 metric tons.  Minimum crew 3.  I have a spreadsheet model already constructed for this purpose,  which I proceeded to run again for these exact numbers.  Masses for the dry weights of capsule and trunk (before modification) are my best guesses,  but their sum matches published data. 

The modification is to install more tanks of NTO-MMH propellants in the trunk,  to a maximum of the 3000 kg quoted cargo capacity for that trunk.  I estimated propellant-tank mass split as 95-5 or a 19:1 ratio,  same as for the landers.  I did not estimate volumes,  although there are 14 cubic meters available in the trunk for this. 

Results That Bound the Design

These are shown in Figure 4.  Payload mass is limited more by the 13.00 ton thrown weight than the 1.35-1.4 km/s delta-vee requirement.  That payload mass is 1760 kg. 

The per person allotment we used for the lander was 200 kg person-plus-suit,  and 100 kg of packed life support supplies.  The life support supplies are probably a bit of an overkill,  so 1760 kg ~ 1800 kg,  and 1800 kg / 300 kg/person is crew = 6 max.  Slightly less actually.  Call it no more than 5 crew at a time,  plus life support supplies,  and no more than about 150 kg of equipment or cargo in the capsule with them,  for the trip to the moon. 

Having the extra delta-vee means we can carry 6 crew,  even 7,   home.  That is a good safety bonus.  Crew Dragon is supposedly rated for the same cargo home as cargo Dragon (3000 kg),  so we are well within that limit. 

This was accomplished by adding 2800 kg propellants to the trunk,  which also adds 147 kg of tank inerts to the trunk inert weight.  That leaves a smidge for any extra plumbing before we hit the 3000 kg limit. 

The only remaining question is for how long a crew Dragon can be parked in lunar orbit before the crew that needs it must come back to it and fly it home.  There are limits to lifetime allowable parked in space.  Perhaps this can be made into some number of months to a year,  given some experience flying the capsule design in Earth orbit.  If a year,  then the lander plan given above is quite feasible just as it is laid out.  If not,  we’ll have to switch out crews on the moon at 6 months,  perhaps.

Figure 4 – Modifying Crew Dragon Into Lunar Orbit Dragon for Falcon-Heavy Launch

Final Remarks

With these two bounding analyses,  I have shown how it is possible to ship 13-ton lunar cargo and crew landers to the lunar orbit with Falcon-Heavy as the launch rocket.  I have also shown how it is possible to ship crews to lunar orbit with the same rocket and a 13 ton modified crew Dragon that has 2.8 extra tons of propellant in its trunk,  connected to the Super Draco thruster systems in the capsule. 

The cargo landers deliver slightly over 5.3 tons to the surface.  The crew landers have a 5.3 ton ascent stage that could carry as many as 6 crew back to lunar orbit. 

At only $17M/delivered ton,  building a practical small experiment station that is permanently occupied becomes easily possible,  at a price far below what was experienced doing the Apollo “flag-and-footprints” stuff during the cold war. 

What makes this feasible is a heavy lift rocket of adequate size to put 13-ton payloads into lunar orbit,  and at a commercial launcher’s far lower price.  This is true flying the rockets fully-expendably.  This capability should become available within the next 1-2 years. 

All that is needed from a vehicle development standpoint is the two versions of the lander designed to these bounding limits,  and then developed and made ready for use.  These share a common descent stage.  That should help lower costs and development time. 

Adding propellant capacity to crewed Dragon with tankage in the trunk is not so much development work,  more of a routine modification that can be tested all-up in Earth orbit,  to make it ready to use.
We’ll need a 2 or 3 seat open electric rover car that weighs no more than a ton.  Between the Apollo rovers and the recent Mars robot rovers,  this should not be a major development item.  

Development,  yes,  just not a “biggie”.  Same for a 2-ton electric front-end loader. 

The hardest nut to crack is a surface habitat that can be assembled from modules that fit within the 5 ton lander payload capacity,  and that can be erected by men on foot in spacesuits with hand tools.  The idea is to assemble it in an excavation done with the front end loader,  and then bury it at least partially with that front end loader. 

This is the kind of thing that could be done within 1 or 2 presidential terms,  which would net returns orders of magnitude greater than Apollo,  for costs orders of magnitude less than Apollo.  

Monday, March 6, 2017

Reverse-Engineered "Dragon" Data

Reverse-Engineering What the Versions of “Dragon” Can Do       
GWJ       2-17-17                 updated 3-5-17

Sources:  Spacex’s website and the Wikipedia articles on cargo Dragon,  crewed Dragon,  and Red Dragon.  There is also DragonLab,  which is a very close variant of cargo Dragon.  These give dry weights for the spacecraft that seem to include the associated trunks,  except in the case of Red Dragon,  which is listed in a very sparse article as “6.5 ton plus payload up to 1 ton”.   Comments made in public by Spacex have indicated the possibility of more than 2 metric tons payload to Mars for some time now. 

Cargo Dragon:  The Wikipedia article lists dry mass as 4200 kg,  and speaks of a chute drop test at 5400 kg that includes a max cargo weight of 2500 kg.  Propellant quantity for the Draco thrusters is no longer on Spacex’s site,  but was once listed as just about 1290 kg.  The capsule has a jettisoned nose cone fairing for ascent,  for which a wild guess is 50 kg. 

The ocean landing test configuration would be capsule dry mass plus max rated cargo,  plus some propellant residual if not jettisoned after entry and chute deployment.  Being toxic,  they should be jettisoned before recovery is attempted by humans.  I assumed zero propellant residuals,  so that the actual capsule and trunk dry masses could be determined in this way:

Both the website and the Wikipedia article list max cargo “up” as 6000 kg,  with at most 3000 kg in the capsule,  and with 3000 kg unpressurized in the trunk.  Max “down” cargo is listed as 2500 kg in the capsule,  with up to 3000 kg of waste in the trunk to be destroyed on entry.  Max cargo available to be carried to the ISS is listed in the Wikipedia article as 3310 kg,  presumably a max of 3000 kg in the capsule,  and the rest in the trunk.   The station’s arm is used to unload items in the trunk.

Weight statements for cargo Dragon can now be estimated to the accuracy that trunk dry mass estimate is accurate,  and that the nose cone mass can be guessed.  For three possible cargo loadouts these are:

Compare the launch weights above with Falcon-9 capability to LEO from Spacex’s website.  If flown as an all-expendable launcher,  the rocket can send 22.8 metric tons to LEO,  and only as a guess probably something close to perhaps 15-17 metric tons to ISS.  All the cargo Dragon estimates shown above fall well within that capability,  at none over 11.5 metric tons.  Whether the booster core is recoverable at 11.5 tons is just not determinable (the website does not list those reduced payload limitations).

One of the things in the weight statement is the set of ignition and burnout weights for the capsule-only,  no trunk.  Cargo Dragon is not operated that way,  however!  It retains the trunk until after the reentry burn.  So it is capsule-plus-trunk ignition and burnout weights that we are really interested in. 

To get those ignition and burnout weights,  you add the capsule-only ignition weight and the total loaded trunk weight for capsule-plus-trunk ignition weight,  from which you delete the propellant for burnout weight.  This leaves out the nose cap,  which was already jettisoned during ascent.  

The cargo Dragon has 18 Draco thrusters arranged in two pods of 4 and two pods of 5 within the outer mold line.  These provide attitude control and maneuvering delta-vee,  plus re-entry delta-vee.  Each Draco is about 90 lb thrust (400 N).  These burn NTO-MMH,  for which one might assume Isp = 335 sec for a “good vacuum” engine design,  meaning a long bell for high expansion ratio.  The corresponding mass ratios (MR) and max theoretical delta-vee capabilities for capsule-plus-trunk are:

Crewed Dragon (Dragon v2):  This is the same basic capsule pressure shell and mold line,  modified for four protruding pods,  each pod containing two Super Draco thruster engines and four Draco thrusters for attitude control and minor maneuver.  The Super Dracos are listed on the Wikipedia article as 16,000 lb thrust (71 KN) each.  They use the same propellants as the Dracos.  Spacex’s website lists the eight total Super Dracos as having 200,000 lb axially-directed thrust (890 KN).  Older versions of the site listed the propellant load as just about 1890 kg.

The Wikipedia article lists dry weight as 6400 kg,  which apparently includes an empty trunk.  This capsule has the same chutes,  a retained reusable nose cap,  crew life support,  crewed interior seats and fitments,  and landing legs.  The trunk is of similar size,  but arranged with conformal surface solar panels instead of folding solar panel wings.  It does have four aerodynamic fin surfaces for aerodynamic stability during crew escape situations.  There is no information available anywhere I can find by which to separate the trunk dry mass from the capsule-plus-trunk dry mass. 

Crewed Dragon operates in space as capsule-plus-trunk,  until after the reentry burn,  when the trunk is jettisoned.  The Wikipedia article lists exactly the same cargo masses and volumes as for cargo dragon.  Unlike cargo Dragon,  crewed Dragon uses the chutes only as a safety backup landing method,  or for landing in the ocean.  Its intended mode is a propulsive landing on dry land with the Super Draco engines,  no chutes at all.  Because of this,  both the capsule-only and capsule-with-trunk max theoretical delta-vees are of interest.  Note however that you cannot achieve both simultaneously,  because there is only one propellant supply to be used for both purposes! 

The best I could do was to simply assume the two trunks were comparable mass in spite of the design differences.  The uncertainty in the resulting data is dominated by that assumption.  Again,  I assumed Isp = 335 sec for an exhaust velocity of 3.285 km/s.  Using cargo Dragon’s trunk mass, the crewed dragon capsule dry mass (which includes the reusable nosecone) is:

This capsule-only dry mass is slightly more than 2 tons higher than cargo Dragon,  but there are the eight Super Draco engines,  an uprated heat shield,  landing legs,  life support,  and crew seats and fitments to consider,  so it is “reasonable”.  From this, one can estimate the same sort of weight statement breakout already reported for cargo Dragon,  including both capsule-only and capsule-plus-trunk ignition and burnout weights.  I did this for only one crew/cargo value,  chosen to approximate a capsule-only delta-vee of 0.7 km/s to compare with Red Dragon.

These figures show comparable values of capsule-plus-trunk delta-vee to cargo Dragon’s ~0.5 km/s,  which is realistic,  considering crewed Dragon is a derivative design,  operating in the same capsule-plus-trunk configuration.  The slightly-higher capsule-only figure is to compare with crewed Dragon’s unmanned derivative Red Dragon (for one-way probes to Mars).  Note that this would vary significantly as crew/cargo is adjusted.  Under the assumptions of 100 kg person,  100 kg suit,  50 kg air and water,  we are talking about 7 crew plus 1050 kg cargo in this 2800 kg loadout.  The weight to launch falls well within the Falcon-9’s LEO capability,  being just about like the heaviest cargo Dragon presented above.

Red Dragon:  This is the crewed Dragon with the crew seats and fitments,  life support,  and chutes removed,  and some equipment racks installed.  The heat shield is reduced in thickness as well.  Since the vehicle is not to be reused,  a jettisonable nose cap like that of cargo Dragon is assumed.  It will need some sort of trunk for launch and for electricity during the trip,  but this is jettisoned before Mars entry.  Course correction burn is assumed trivial,  so that essentially the entire propellant load is available for powered landing on Mars. 

There are no available data for the masses of any of the change items just discussed.  All that is available are wild guesses and educated guesses.  The lighter heat shield I estimated as a reduction from 8 cm thick to 6 cm thick,  on a flat circle 3.7 m diameter,  and a specific gravity of ~0.3 for PICA-X.  I just rounded off to the nearest 10 kg.  It’s just too rough not to round off like that.    

I have just assumed the same trunk mass as I used for cargo and crewed Dragons.  Trunk mass dominates the uncertainty,  being the largest item.  I simply took the crewed Dragon estimated dry mass,  and subtracted things.  Those guesses are listed in this estimate for Red Dragon dry mass:

If I load this vehicle with 1 ton of cargo and 1890 kg propellant,  then mass at entry is 7640 kg,  which is not far at all from the 7500 kg indicated the Wikipedia article!  1 ton of cargo is what is indicated as deliverable to Mars in that same article.  (Some public announcements indicate that 2-4 tons are actually under consideration at Spacex.)

Using these figures,  the weight statement for Red Dragon can be roughly estimated.  What is of interest here is the capsule-only delta-vee,  as a function of cargo delivered to the surface of Mars.  Bear in mind that an utter minimum delta-vee capability for powered landing will be near 0.7 km/s,  the Mach 3 point coming out of atmospheric entry hypersonics.  There’s very little in the way of gravity and drag losses to correct the theoretical delta-vee in this scenario.  The error is less than the uncertainty in the basic requirement.

The 0.7 km/s figure is pretty rough,  that being 3 times the nominal speed of sound in the Martian atmosphere at something like 5 km altitudes.  This could vary quite a bit.  In order to successfully land reliably,  you actually need a little more delta-vee to cover final maneuvering around obstacles. 

The Mach 3 point is also a bit arbitrary,  that being only the definition of min-hypersonic for blunt objects.  Prior probes deployed chutes at local Mach 2 to 2.5,  although they did this much higher up (15-25 km altitudes or more).  Waiting to lower speeds lets you penetrate to lower altitudes,  while heavier items also penetrate to lower altitudes,  simply because of higher ballistic coefficients. 

The 0.7 km/s “requirement” I use here is thus just a figure of merit,  although it is actually in the ballpark of the true requirement.  

These numbers are too rough to judge “for sure”, but it looks like 1 or 2 metric tons should be easily deliverable to Mars with Red Dragon,  just like Spacex has indicated.  These numbers say 3.2 tons is getting to be quite marginal,  but that would actually depend upon what the true landing delta-vee requirement really is.  Note that the requirement would vary with location and season across Mars,  as that atmosphere is much more variable in its density than is Earth’s. 

Spacex’s website lists Falcon-Heavy as able to send 13.6 metric tons to Mars,  flown fully-expendably.  All these configurations fall within that capability.  Even at 4 tons cargo,  the weight to launch would be just about 12 tons,  which still falls within the launch capability.  Therefore,  it will be landing delta-vee that sets the payload deliverable to Mars!  That may explain the “extra propellant” remark found in the Wikipedia article.  Whether 11 tons at launch is small enough to recover the booster cores is unknown. 

DragonLab:  this is cargo Dragon with an instrument bay between the pressure shell and the outer mold line.  The door on this bay opens to space,  and recloses before entry.  It is otherwise the same as cargo Dragon,  so no separate analysis is done here.    Use my figures for cargo Dragon to represent DragonLab.

Using variants of Red Dragon as unmanned one-way probes elsewhere:  This depends on the delta-vee requirement to land,  relative to the vehicle capability.  For airless destinations with direct landings from the interplanetary trajectory,  several-to-many percent above the body’s escape speed is a figure of merit.  Example:  Mars escape 5 km/s,  typical direct entry speed 6 to 7 km/s.  Call it 1.5 Vesc as a typical figure of merit.  Remember that this is only a very crude estimate,  unlikely to be correct!

Potential destinations include the moons of Jupiter and Saturn,  the asteroids,  and our own moon.  Mars and Venus have atmospheres that allow aerobraking to a propulsive landing,  as does Titan at Saturn.  Mercury requires all-propulsive landing,  as does the moon.  Values of escape speed and surface gravity strength follow.  I did not include Venus because landed vehicle lifetime would be too short,  if it made it down to the surface at all.  I also did not include Earth itself.

The first thing apparent from the list is that any of the asteroids,  Titan,  and Mars all seem to be within reach of Red Dragon on a Falcon-Heavy,  just as it is.  The significant atmospheres of Mars and Titan make aerobraking feasible,  the rest are airless or so tenuous as to make aerobraking infeasible.  Like the Earth’s moon,  the moons of Jupiter seem out-of-reach,  due to escape speeds that are too high.

As a one-way probe destination,  Earth’s moon is interesting on its own.  Key here is getting into lunar orbit using the upper stage of the launch rocket,  without using any of the Dragon’s propellant.  As it turns out,  that delta-vee requirement for the launch rocket (no more than 12.4 km/s) is very similar to that for sending things to Mars (at least 12.1 km/s).   Those figures include 5% gravity/drag losses on the first 8 km/s of that delta-vee.   From there it takes 1.68 km/s to make a powered landing.  That’s out-of-reach for Red Dragon without considerable extra propellant. 

It might be more desirable to instead enter lunar orbit with a crewed Dragon,  and let them rendezvous in lunar orbit with a separately-sent lunar lander.    

The lander must descend and ascend to a delta-vee of essentially the orbit velocity each way, for an utter minimum of something like 3.36 km/s plus a tad for gravity losses.  At the moon,  the crew needs a minimal place to ride,  not a full capsule,  but the stage does need landing legs.  It needs to carry surface stay gear and a rover,  as well.  That design is not explored here.

As for the missions to the outer moons,  there needs to be a fairly-large propulsion stage added to the Red Dragon.  It seems like the Dragon probe assembly could be sent to Earth orbit on a Falcon-9,  and the propulsion stage sent there with a Falcon-Heavy,  to be docked together in orbit,  and launched on its mission from there.  Very much better information on velocity requirements is needed to size such an exploration stage design.  That is not addressed here. 

Conclusions:  Red Dragon as presently envisioned works for Mars,  Titan,  and any of the asteroids.  The other outer planet moons require a fairly large powered stage added to the Red Dragon to achieve the necessary delta-vee for capture and landing.  The combined weight exceeds Falcon-Heavy capabilities for direct interplanetary trajectories,  so that something other than direct launch to interplanetary travel is required.  Falcon-Heavy is able to fling 13.6 metric tons to Mars,  perhaps 12-13 tons into lunar orbit.

For Apollo-like lunar missions,  crewed Dragon with extra propellant in the trunk (yielding near 1.6 km/s capability) fits one Falcon-Heavy,  and a lander not based on a Dragon fits another Falcon-Heavy.  These need to weigh under 12-13 metric tons to be successfully launched direct from Earth’s surface,  and must rendezvous in lunar orbit.   Red Dragon itself,  as it currently is envisioned,  seems unattractive for one-way missions to the moon,  with a direct landing delta-vee near 2.4 km/s.  However,  whatever added propulsive stage works for the outer planet moons would work at Earth’s moon. 

Addendum 3-4-17:

I looked very crudely at how much propellant to carry in the trunk to enable a crew Dragon to depart from lunar orbit and still have propulsive landing capability at Earth.  This would be for a crew of only two,  with their suits,  and about 500 kg of supplies and samples.  I did actually add propellant tank inert mass to the trunk (about 147 kg),  which requires iterations.  One needs the trunk for electrical power during the trip home to Earth,  so it is capsule-plus-trunk mass ratio and delta-vee that are pertinent. 

Results are given in the Figure just below.  It is slightly over 12 metric tons as launched.  Falcon-Heavy just might be able to deliver this to lunar orbit with a delta-vee of no more than 12.4 km/s,  because Spacex’s website says it can send 13.6 metric tons to Mars (with an estimated delta vee no less than 12.1 km/s).  Those delta vee estimates are my calculations made using Hohmann min-energy transfer ellipses at orbital semi-major axes that produce the largest delta-vee requirements,  and for which I applied 5% gravity and drag loss to the first 8 km/s getting off the surface of the Earth. 

The capsule-plus-trunk needs about 0.8 km/s sec to depart lunar orbit onto a free-fall homeward trajectory to an aerobraking entry.  It perhaps needs about 0.7 km/s capability to cover a propulsive landing on land,  per the basic design of the crewed Dragon.  The total is 1.5 km/s,  for which a few percent margin gets you quickly to 1.6 km/s capability.  The configuration shown in the figure has such capability by adding some 2800 kg pf propellants to the trunk,  in tankage weighing about 147 kg.  That total is 2947 kg,  within the 3000 kg cargo rating of the trunk.  It is the capsule that is lightly loaded at only 2 suited persons plus modest supplies.  Adding a third crew person pushes the total closer to 13 metric tons,  just about the limit that Falcon Heavy could possibly deliver to lunar orbit.  

Addendum 3-5-17:

The same sort of added propellant in the trunk could also be done with Red Dragon.  Note that I have already pretty much defined the maximum propellant already at about 2800 kg.  Red Dragon being similar to crew Dragon,  the capsule-plus-trunk delta-vees will fall far short of what is needed to capture and land on the outer planet moons (1.6 km/s or thereabouts versus 2.5+ km/s).  The capsule-only value indicated in the figure is wrong,  because most of that propellant must be in the trunk.  

I’m not at all convinced that a capsule is the best vehicle by which to send probes to the airless bodies,  including the asteroids,  which Red Dragon could certainly reach.  But Red Dragon should serve very well for probes to Mars,  and maybe Titan.  With aerobraking entry,  a capsule and heat shield are necessary. 

Crewed Dragon could easily become part of a system to send men back to the moon with very few changes from what will start flying this year.  Flyby missions need no changes,  and orbit missions require some extra propellant in the trunk for the Super Dracos.  Landings will require a separate lander sent ahead to lunar orbit unmanned,  for the crew to rendezvous with and utilize.  If such a lander totals under 12-13 metric tons,  Falcon-Heavy could fling it to lunar orbit launched directly from Earth’s surface. 

Final Comments as of Posting 3-6-17:

Enjoy!  These figures may not be exactly right,  but they are pretty close.  I suggest you create spreadsheets to calculate delta-vee in capsule-plus-trunk configurations for cargo Dragon and for crewed Dragon,  and for both capsule-plus-trunk and capsule-only Red Dragon.

Then,  given real delta-vee data to reach a destination,  you can compute for yourself whether the corresponding Dragon configuration can reach it.  Be sure to factor-up astronomical values for gravity and drag losses where those apply.  This is required before you size mass ratios and weight statements.

There's no information out there about it,  but I would hazard the guess that Spacex is already looking at versions of crew Dragon and Red Dragon that have extra propellants in the trunk.  That just makes too much sense for them not to be doing that.