Death for Hasan

I just heard on the radio that Nidal Hasan got the death penalty for his military court martial conviction as the Fort Hood shooter.  He very most definitely deserves it.  I wish it could be carried out sooner,  rather than later.  And it ought to be hanging-without-the-drop,  the old way. 

Hasan was tried on workplace-violence charges instead of terrorism charges,  because there is nothing in the Uniform Code of Military Justice to cover terrorism by a soldier.  There should be.  The bureaucrats who use this legal shortfall as an excuse not to provide aid to Hasan's victims,  should be tarred and feathered,  as bad public officials often were,  in my grandfather's day. 

The officers who knew that Hasan had loyalty and fitness problems should be court-martialed along with him,  for dereliction of their duties.  They should get hard labor,  at the very least. 

Why Hasan was not tried for treason,  I do not understand.  Out of his own mouth,  he said "he joined the other side",  meaning specifically the enemies we are currently at war with.  More than the required 6 persons heard this.  He said he shot those people at Fort Hood to protect the Taliban and others like them,  the very folks we fight in Afghanistan.  More than the required 6 persons witnessed that statement,  too.  Killing Americans,  to provide aid to the enemy Taliban,  meets one of the definitions of treason.  And that is in the Uniform Code of Military Justice.

Somewhere,  somehow,  we (military and civilian) have become too "eaten-up" in political correctness to face the hard truths of our times,  and deal with them.  This includes dealing-out real justice when it is needed.  Real justice doesn't lie in precedents,  and most certainly not in political correctness,  nor in polls.  It lies in good old common sense and a solid upbringing in ethics.

This Hasan was a home-grown,  self-radicalized,  terrorist and traitor.  He converted himself to the belief/ideology that he must "kill for God",  as did the would-be Fort Hood bomber,  Naser Jason Abdo.  Those two Tsarnaev brothers who bombed the Boston marathon were immigrants,  but they self-radicalized over here.  Same radical belief system/ideology. 

I don't know the relative education level of the other three,  but Hasan was a psychiatrist with a terminal college degree.  I don't understand how an educated person could fall for that radical Islamist propaganda that wants its adherents to "kill for God",  but he did,  and so have many others. 

Any radical/fundamental group that wants you to "kill for God" is very most definitely not speaking for the God that I know.  Islam is not the only faith afflicted with vicious radical subgroups like this,  they all have them.  It's just that the Third World,  which is pretty much synonymous with the Muslim World,  is in intense turmoil in our times.  Radical Islamists are more visible simply because there are so many people of all types involved in all that turmoil.  And it spills over here. 

Generally speaking,  I am totally against "ratting people out".  But for self-radicalizing terrorists like Hasan and the rest,  I make an exception.  So should we all.   Everybody should be watching for this.  Otherwise we'll have another 9-11 event,  and it'll be some of our own who did it. 

On Egypt’s Troubles

Update 4-9-17:  It would appear that the military takeover failed to return Egypt to its people,  instead becoming just another military takeover.  The "good" they did was ending the impending Iran-style religious dictatorship. The "bad" was just substituting another style of dictatorship.  Very disappointing.  

Original Article:

I saw the first fully-truthful news story about the troubles in Egypt in the Wednesday 8-21-13 Waco “Trib”,  on page 6a,  sort-of hidden on the backside of section A.  That’s the first time I have seen this truth printed,  and I have not yet seen it anywhere on television or the internet. 

Update 9-11-13:  With the Syrian situation breaking toward a possible divestment of chemical weapons,  brokered by Russia,  I have begun to see some hints of the real truth showing up in a few of the TV news stories. 

Kudos to the “Trib” for running it.  This was an Associated Press story from Cairo,  written by one Hamza Hendawi,  obviously a local.  It is brief,  but contains all the information previously seen only as unconnected tidbits,  here and there. 

Morsi is Muslim Brotherhood to his core,  and was in prison as such,  when the revolution that overthrew Mubarak occurred.  The story brings up the connection between Morsi and Iran’s proxy army Hamas,  a connection that allowed his escape. 

The Muslim Brotherhood is a rather extreme-fundamental Islamist group,  not unlike the Taliban and some others.  It has been the “mother ship” for many of these other Islamist organizations,  according to the story.  These people are quite intolerant of personal freedom.

Morsi got elected Egypt’s president,  mainly because he was “not-Mubarak”.  The people passed-over his opponent in their first free election,  because of his taint of prior public service under Mubarak.  Morsi created,  by his presidential appointments,  a network of Muslim Brotherhood figures who would go along with the Brotherhood agenda:  to take over control and establish a strict religious dictatorship.

This group even modified the new constitution to aid their takeover,  and put that up for public approval.  As it became apparent that Egypt was headed for a fundamentalist dictatorship,  public opposition mounted.  When this became widespread massive public demonstrations,  Morsi’s government responded (not surprisingly) with violence. 

That’s when the Egyptian army stepped in,  on the side of the people,  and removed Morsi in a quick coup.  There have been very violent pro-Morsi demonstrations ever since,  but the anti-Morsi turnout has actually been larger (something not often reported). 

Yeah,  it was a coup.  Yeah,  the army is in control,  and has responded in-kind with violence,  to that violence perpetrated by the Morsi supporters.  We have yet to see the army stage a free election,  but I think they eventually will,  and the Egyptian people will be the better for it.  The army saved them from a harsh religious dictatorship that would have resembled the Taliban government in Afghanistan. 

According to the article,  the thinking in Egypt is that the Muslim Brotherhood is finished.  Either they ally themselves with Al Qaeda-like terrorists,  or else they go into hiding for a very long time.   It looks to Mr. Hendawi (and to me) like the new government,  whatever it turns out to be,  will have no place for the Muslim Brotherhood.  And that’s a good thing.

Here’s my take on it:  what you have really witnessed in Egypt is a cultural civil war between those wanting a religious dictatorship,  and those opposed to it.  Those two groups will temporarily ally,  in order to overthrow secular dictators (or foreigners -- update 9-11-13).  That throwing out dictators and having a civil war afterward is what the so-called “Arab spring” is really all about. 

Generalizing,  you have witnessed this same civil warfare in Iraq (it’s still going on,  even though we left),  Afghanistan,  Pakistan,  Libya,  and several other Middle Eastern countries.  It really doesn't matter whether you call them Al Qaeda,  Taliban,  or a whole host of other names,  it's always between those who want a religious dictatorship vs those who do not.  ---  9-11-13 update. 

Syria is still trying to get rid of a secular dictator (Assad).  We dallied around too long,  before deciding to help them.  Now,  the Syrian opposition has swelled with extremist foreign fighters,   who will start its cultural civil war as soon as Assad is overthrown.  They are already beginning to attack their opposition brethren who do not share their wish for an Islamist dictatorship.

Iran had its revolutionary overthrow of a dictator (the Shah,  that we put there,  which is why they hate us) long ago.  Their cultural civil war never erupted at that time,  they just went straight for the religious dictator (Ayatollah Khomeini) because he was “not-the-Shah”.  There was a hint of a revolt against the religious dictatorship recently,  but it aborted (we failed to help them). 

The bad news is that the fraction of local populations who want extremist religious dictatorships is so high.  The good news is that they are still minorities in some important places,  like Egypt.  I wish our State Department and our CIA understood this fact-of-life better.  Our track record dealing with this region over the last half a century is very poor. 

Update 9-24-13:  Recent AP news stories have described the dismantling of the Muslim Brotherhood's support network.  This was a network of needed social services not provided by others,  which then served as a "fundamentalist pulpit" from which to recruit extremists. 

It would appear that in Egypt the civil war is being won by those who do not want a religious dictatorship.  The question now is will they learn the lesson of the needed services?  If not,  those who want a religious dictatorship will just slowly rise again. 

Applying Ramjet to Launch Accelerators

I have seen (and been asked about) using constant "averages" for estimating ramjet performance.  This would be for launch applications,  either vertical assist,  or for horizontal launch.  I also need  to update readers as to the status of my ramjet book (see the "Ramjet Cycle Analysis" posting dated 12-21-12 below). 

Well,  I think arguments based on mass ratio and “average” Isp are too crude to get you anything useful for ramjet (or any other airbreather).  You need a real cycle analysis,  which should be a subroutine in a real trajectory code,  or which you can use for point performance calculations over a flight envelope for empirical correlation.  I’ve done both,  they’re both effective approaches to first order.  Fixed averages are not.  Sorry,  that’s a simple fact-of-life.

You do need to understand thrust and drag accounting,  because if you don’t,  it is really easy to leave out some very important drag forces in your force balance.  I’m not talking about basic ram drag here (the “airbreather’s burden”),  I’m talking about things like additive drag,  spillage drag,  diverter drag,  and bleed drag.  These are quite important,  both at takeover,  and at very high flight speeds.  These are neither trivial to understand,  nor trivial in their effects.  You need some training in propulsion aerodynamics for these.  This isn’t basic physics textbook stuff,  and never will be. 

All this stuff will be in my ramjet book,  which is not yet ready for publication.  Its intended audience is engineers working in ramjet propulsion,  whether for missiles,  or for launch vehicle work.  I’m still trying to “rough-write”-down all the topics,  but I think I have most of them documented in rough form,  just not all of them.  All this stuff is currently very rough first-draft stuff,  and will need extensive re-organization and re-write,  before it is book-ready.  But,  I really am working on it. 

There are two speed ranges for ramjet design,  “low” and “high”.  Low speed range designs have simple pitot (normal-shock) inlets,  convergent-only nozzles,  and can be ignited at subsonic speeds.  They will show nacelle thrust greater than drag down to very low speeds,  but will have specific impulse lower than composite solid rocket,  below about half a Mach number.  Peak specific impulse potential is at about Mach 1.1 or so,  at about half or 2/3 the max Isp potential of supersonic designs.  Max useful speed is about Mach 2,  or maybe Mach 2.5 at the very outside.  With hydrocarbon fuels of almost all types,  about the biggest nozzle throat/combustor area ratio is 0.65,  limited by flame-holding considerations.  Performance at lower area ratios is inherently lower. 

High speed-range designs feature external compression features like ramps or spikes that protrude ahead of the inlet cowl lip.  They also have almost-zero thrust potential below about Mach 1.6 to 2.  But,  they work just fine to about Mach 5-or-6,  depending far more on vehicle drag characteristics,  than anything about the ramjet engine design.  With kerosene fuels,  peak Isp potential is around 1200-1300 sec at about Mach 2.5-ish,  lower slower,  and lower faster.  Nozzles are C-D,  but exit “bell” area ratios are closer to 1.5-max,  than anything to do with the expansion ratios one sees in rockets.  With hydrocarbon fuels of almost all types,  about the biggest nozzle throat/combustor area ratio is 0.65,  limited by flame-holding considerations.  Performance at lower area ratios is inherently lower. 

These things can be very lightweight,  depending upon whether it has to be re-usable or not.  The “best” designs have been one-shot missile designs,  with an ablative combustor liner,  for missile speeds up to about Mach 4.  External heat protection is also an issue,  from about Mach 3 on up for reusable designs,  even with steel construction.  There are air-cooled perforated liner designs from the 1940’s and 1950’s that would actually work to Mach 6 on a transient,  exclusive of external heat protection problems.  There are ablatives that would work externally to Mach 6 on a transient,  but these have replacement issues.  Missiles generally always use ablatives inside,  and maybe outside,  if needed.

There is my oddball ceramic-ceramic composite combustor liner material,  which offers considerable potential for a re-usable combustor.  It might also serve as external heat protection,  for a fully-re-usable design.   This is still an experimental material,  though.  (See also the 3-18-13 posting "Low Density Non-Ablative Ceramic Heat Shields" below).

Ramjets require boosters to reach takeover speed:  about Mach 0.5 to 0.8 for “low speed” designs,  and about Mach 1.6-to-2 for “high speed” designs.  For one-shot missile applications,  the best choice has proven to be the “integral rocket-ramjet”,  wherein a solid rocket booster is cast or loaded within the ramjet combustor.  This requires an appropriate ejectable booster nozzle nested within the ramjet nozzle,  and some sort of inlet duct obturator,  usually ejectable or frangible port covers.  Re-usable launch applications might well be “best” with parallel-burn rocket and ramjet engines in the same airframe.  It really helps if the rocket and the ramjet use a common fuel. 

From a flame-holding standpoint,  I think the dump combustor has “way-to-hell-and-gone” more potential than the V-gutter,  or can,  or “colander” (or any other type of obstruction-type) flameholder.  Dump combustors have very little sensitivity to dump plane speeds,  compared to any of the blockage-element types.  Variable speeds at the dump are inherent with launch accelerators,  whether vertical-launch or horizontal takeoff.  Almost no textbooks describe dump combustors.  My book will. 

Ramjet liquid fuels can be any kerosene (or kerosene-like synthetic),  or any liquifiable hydrocarbon.  The early engines with subsonic ignition used mainly low-grade gasoline.  Today,  in supersonic-inlet designs,  JP-4,  JP-5,  JP-7,  Jet-A,  Jet-B,  Jet-A1,  RP-1,  K-1 kerosene,  a synthetic variously known as RJ-5 or Shelldyne-H,  and even liquefied methane,  are all very attractive candidates. 

I have even used propane,  but it and LPG are not all that attractive,  for their inherently-heavy fuel storage considerations.  LCH4 will require extra care to insure full vaporization,  and extra care with flame-holding issues.  RJ-5 is a synthetic that resembles kerosene,  except that its density is substantially higher.  It was used in ASALM-PTV,  with one test that reached Mach 6. 

I hope the book might be available in a year or two.  It’ll be the ramjet analog to the famous (or infamous) “drag bible” written long ago by Hoerner. 

GW

update 8-23-13:

The "high speed range" ramjet designs would be most applicable to a horizontal-takeoff (HTO) two-stage design featuring a winged airplane as its first stage.  The most important consideration for selecting the best staging point seems to be as fast as possible.  The upper speed limit for the ramjet airbreather is about Mach 5 to 6. 

This is limited more by the vehicle drag,  than anything to do with the ramjet design,  although the minimalist variable inlet geometry of constant shock-on-lip seems to be the best enhancement one could try.  Max specific impulse (Isp) potential with hydrocarbon fuels is about 1300 sec near Mach 2-to-2.5,  and about half that,  at the upper and lower ends of the speed range. 

Scramjet (supersonic combustion ramjet) might fly much faster,  but has a far higher takeover velocity (Mach 4,  1.2 km/s),  and (worst of all) is simply not technologically-ready for application.  The inlet,  combustor,  and nozzle geometries for scramjet are completely incompatible with those of the ramjet.  If included at all,  it would have to be yet-a-third engine type carried on the first stage.  That tends to increase both stage inert weights and vehicle drag.  Heat protection is also an extreme problem above Mach 6,  especially for shock-impingement zones. 

As a second-most important consideration,  this ramjet staging point is deeper in the atmosphere than most people would assume:  only about 60,000 feet (18.3 km).  The frontal thrust density of the ramjet depends very strongly upon ambient air pressure,  and it simply takes too long to accelerate if the air is any thinner than that.  As it is,  there is no thrust margin-over-drag at max speed to support a climb,  without serious and sudden deceleration.  That means rocket thrust from somewhere must be added,  to support a short pull-up maneuver transient at staging. 

(As an aside,  scramjet suffers from almost exactly the very same thin-air altitude limitations.)

The third-most important consideration at ramjet staging is the pull-up trajectory path angle.  Something around 40 degrees above horizontal seems to be about right.  This relieves the second stage of any lift (or major thrust vector) capability required to pull up.  It may simply fly a ballistic gravity turn trajectory from staging.  This is a serious consideration,  since the delta-vee (mass ratio) required of the second stage is quite significant.  That's because Mach 5 to 6 at 60,000 feet (18.3 km) is only about 1.5 to 1.8 km/s velocities.  A total of 7.7 km/s plus gravity and drag losses is required to orbit. 

This leads one to a combined rocket/ramjet winged HTO aircraft as the first stage.  It could carry either a rocket pod or a rocket airplane as the second stage.  The characteristics of this first stage aircraft,  if a truly reusable design,  would resemble more a supersonic bomber,  than any of the prior space launch vehicles we have ever flown,  including the shuttle.  There is a practical size limit,  which makes this (most likely) a small payload niche application,  probably around 5 tons max. 

It would take off in rocket power,  accelerate on rocket to takeover speed,  then climb on ramjet.  Once at altitude,  it would pull over level and accelerate in ramjet to stage speed.  With some rocket help,  the vehicle pulls up sharply for second stage release.  Then it cuts off rockets and throttles-back its ramjets,  and returns to launch site in ramjet thrust at low supersonic speed (for best range).  It's a dead-stick glider for landing,  except for enough rocket propellant reserve to support a "go-around". 

Update 9-12-13:  new concept I hadn't considered before.  One could pull up in parallel burn,  but not stage yet.  Transition back to all-rocket,  and climb to staging at a little bit a higher speed and altitude.  The first stage airplane is a bit bigger because it has to contain more rocket propellants.  But,  the faster and higher the stage point (at steep path angle),  the lower the velocity requirement imposed on the second stage,  and the bigger the payload it can carry to orbit.  It'll be some sort of tradeoff of payload vs first stage size,  probably constraint-limited by the square-cube law scaling "landing gear" problem at launch weight. 

I don't yet know whether ramjet-assist to a vertical-takeoff (VTO) launch rocket would actually be worthwhile.  But for the VTO rockets we are accustomed to designing,  the vehicle leaves the sensible atmosphere (max 80,000 feet,  24.3 km) at speeds near Mach 2 (only about 0.6 km/s).  So,  I am sure the "low speed range" ramjet design is best suited,  and should be staged off for recovery at that thin-air point,  long before the first stage rocket core burns out (typically well outside the atmosphere at about 3 km/s). 

This kind of ramjet provides useful subsonic thrust from about 0.7 Mach (about 0.2 km/s) up to the low supersonic speed at staging.  (Mach 2 is about the max useful speed anyway.)  Max Isp potential of this kind of design is about half to 2/3 that of the supersonic types,  near Mach 1.1-to-1.2,  and half or less of that,  at the slow and fast limits.  I rather doubt that such ramjet strap-on pods would ever exceed about 25% of the thrust at low altitudes,  far less as the staging point is approached,  but I could be wrong,  as I have not yet fully researched that kind of design.

But,  if this is actually attractive,  the way to make it reusable is very definitely the strap-on pod approach.  Even with ballistic fall-back,  recovery will be very near the launch site.  I rather suspect that some kind of folding wings and fins would turn the strap-on into a big remote control aircraft,  that could be runway-landed,  on land adjacent to the launch site.  The logistics of that offer very low recovery and refurbishment costs. 

VTO rockets are always short on takeoff thrust.  The integral booster approach,  one-shot as it is,  might well actually be very attractive,  as a takeoff thrust enhancement available from the ramjet strap-on pod.  This does put some limits upon the internal combustor heat protection scheme,  since solid rocket pressures are quite high.  Ablatives may be the only practical answer. 

For such strap-on pod designs,  it would be well to separate the combustor/booster case from the tankage and inlet hardware.  These cases might (or might not) be refurbished and reused,  while the rest of the hardware definitely could be easily reused. 

Update 9-12-13:  Another thought for the two stage airplane scenario would be to solve the square-cube law "landing gear" problem by going to vertical launch,  then bending over to the same flattish ramjet acceleration to Mach 6,  before pulling up again to stage.  That's a thrust-enhanced turn deep in the atmosphere:  gravity and drag losses are simply enormous.  Plus,  to get the far larger takeoff thrust,  it'll drive you toward integral solid rocket boosters inside the ramjet engines,  a major limitation on designing for reusability.  So I don't recommend going that way,  for technical reasons,  not to mention the psychology.

The psychology has to do with traditional rocket launch-type thinking versus traditional aircraft-type dispatch thinking.  Vertical rocket launch,  especially with one-shot components like integral boosters in the ramjets,  leads to designs that have enormous logistical support tails.  In contrast,  thinking like an airplane leads one toward very low logistical support,  and thus very much lower costs.  This "high-cost rocket launch logistics thing" has been true of government designs since the end of WW2.  SpaceX and ULA commercial launch rockets with reduced logistics that reduce cost are the recent exceptions that actually prove the rule. 

Better to look like an airplane so you think like an airplane.  You're far more likely to get to a lower launch cost that way.

GW

Entry Issues

This posting concerns entry dynamics and heat protection.  There are two regimes,  crudely separated at speeds of 10 or 11 km/s.  Below,  convective heating dominates,  and to zeroth/first order,  that’s all you need to consider.  The stagnation point has a crude,  easy convective heating correlation that gets you into the ballpark.  The other surfaces have other,  still-empirical models that are less familiar and not so easy to use.  But afterbody heating is less severe than stagnation,  often far less. 

Above that speed,  you must consider radiative heating from the plasma sheath surrounding the spacecraft,  and source conditions vary all around it.  None of the models for this are simple at all. 

 

Coming back from Earth orbit,  you are moving in the vicinity of 7.7 km/s at atmospheric interface,  about 135 km altitude.  You are also moving at a very shallow trajectory angle,  unless you are very wasteful of retro thrust fuel.  Because you are moving at less than escape speed,  you cannot bounce off into space,  although you might skip unexpectedly far downrange. 

Convective stagnation point heat rate per unit area is proportional to the square root of ambient density,  inversely proportional to the square root of the “nose radius” facing the flow,  and proportional to the cube of the speed.  That equation is entirely empirical,  and dimensionally-inconsistent,  but it does work quite well to first order.  The nose radius dependence is why space capsules have blunt heat shields:  the blunter,  the lower the peak heating rate at stagnation.  The effect is quite dramatic. 

You can find this heating correlation in a variety of references.  The most recent is the Justus and Braun EDL paper,  but I had to chase this back to references from the 1950’s and 1960’s before I found a reliable value for the constant of proportionality. 

There is also a simplified entry ballistics analysis presented in Justus & Braun,  that traces back to Julian Allen at NACA in the 1950’s.  I’ve been using it as a zeroth/first order ballpark design model.  I had to correct the heat transfer items in it,  but not the basic dynamics items,  before I could use it. 

The gas/plasma total temperature around the afterbody,  or anywhere it has been shocked-down to local subsonic,  is very crudely numerically equal in degrees K to the vehicle velocity in m/s.  This is another empirical approximation,  and it is not strictly correct,  but it is in the ballpark.  It reflects kinetic energy going into ionization instead of internal energy (temperature). 

The drag coefficients of blunt objects are crudely constant over the range from Mach 5-ish to Mach 25+.  For a capsule shape,  there is a blockage or frontal area associated with the shape that is used as the reference for the drag coefficient.  The same area is used in ballistic coefficient:  W/CD*A

There were 3 original heat shield concepts in the 1950’s:  heat sinks,  ablatives,  and re-radiative (or refractory) concepts.  Heat sinks were then,  and still are today,  a massively-heavy (and therefore undesirable) solution.  The refractories back then were also very heavy (usually tungsten or super-alloy metals and/or monolithic chunks of graphite,  or metal-graphite combined),  leaving ablatives (as heavy as they were back then) as the lightest-weight and most practical solution.  That’s why all the early manned capsules had silica-phenolic (or something closely related) for their ablative heat shields. 

Today,  we have low-density ceramic refractories (shuttle tile),  although these have much lower surface temperature limits than the ablatives,  and,  they are fragile.  We also have lower-density ablatives,  most notably PICA-X.  (And there is my oddball experimental material,  which is much tougher than shuttle tile but not quite as lightweight,  although lighter than PICA-X.)

There are also sacrificial-liquid-coolant schemes.   These are somewhere between heat sinks and ablatives,  and inherently tend to be quite heavy.  This weight is driven more by the integrated total heat absorbed (which sets the coolant mass to be expended),  than just the peak heating rate (although that sizes the coolant flow rate). 

Analysis of surface temperature and survivability of ablatives is a tough calculation problem,  driven by empirical models.  Although,  with the old phenolics,  surface temperature typically fell near 3000 F (1920 K,  1650 C).  Refractories are a little easier,  if they are low-density,  as the conduction pathway is cut off,  completely unlike the old metallic and graphite refractories.  The re-radiated heat simply has to balance the convective input. 

What that means is that we can estimate the surface temperature of the low-density refractory,  given an estimate of its spectrally-averaged emissivity,  and a value for the convective input.  I did this for low-density ceramics at peak stagnation convective-input values.   I found that low ballistic coefficient shapes under 200 kg/sq.m with very blunt heat shields (and I do mean nearly flat) can reduce peak stagnation heating to 25 W/sq.cm,  for entry from LEO,  far lower at Mars. 

Given a black surface average emissivity of 0.8,  the peak skin temperature from that heat balance is about 1290 C returning from LEO.  That’s cool enough for the ceramic to survive entry from LEO,  even at the stagnation point,  unlike the application on shuttle.  The failure mode above that skin temperature is not melting,  but solid phase change-induced shrinkage cracking. 

Capsule shapes have a blunt heat shield,  and an afterbody shape that is pretty much arbitrary,  except that it must be aerodynamically stable with the cg position.  Mercury,  Gemini,  and Apollo all had nose radius/diameter ratios not very far from 1.0-1.1. 

If the afterbody is more-or-less conical,  you can “fly”  at angle-of-attack to the slipstream and generate some lift for trajectory control during entry,  without angling that afterbody into the main slipstream.  All this takes is attitude thruster fuel,  and not very much of it.  This is a well-known and well-proven technique,  dating back to Gemini in the 1960’s. 

The afterbody is more-or-less immersed in more-or-less transonic plasma,  which is quite hot.  Conditions at entry interface 7.7 km/s would be in the ballpark of 7700 K,  just at extreme low density.  As the speed decreases,  plasma temperatures fall,  but density rises due to the descent.  This heat transfer environment is far less severe than that on the forward-facing heat shield. 

That is why sheet metal heat-sinking was adequate for afterbody structures on Mercury and Gemini.  They had corrugated metal shell panels without any ablative at all.  But,  had the capsule tumbled,  it would have been destroyed.  That’s why Apollo had ablatives on its afterbody,  along with the demands of the faster,  hotter entry coming back from the moon. 

Coming back from the moon,  speed is very nearly Earth escape at 11 km/s.  Radiative plasma effects are becoming important,  and there is the definite possibility of bouncing off the atmosphere into deep space,  if the trajectory is too shallow.  Convective heating rates are way far higher,  ruling out ceramics at the stagnation point,  leaving ablatives as the only practical choice for stagnation regions.  About 2 degrees from horizontal is what Apollo used.  Steeper is too much heating at too many deceleration gees.  This is precision trajectory control during the moon-Earth transit,  no way around that. 

Once into entry,  there is less need for trajectory control,  other than controlling attitude with small thrusters,  unless a precision landing point is needed.  Apollo used the same angle-of-attack trajectory control during entry as Gemini,  and it was quite successful.  Circular error probable was around 1 or 2 miles (1.6-3.2 km). 

Depending upon the nature of the cargo to be brought from the moon,  no afterbody shell may be needed at all.  Bulk metals or minerals,  for example,  can just heat sink their way through 3 minute’s exposure to transonic plasma decreasing from about 11,000 K effective at interface. 

A design like that is just a cargo deck with heat shield on one side,  and the cargo plus a guidance package on the other.  Other types of cargo may require protection from the plasma sheath (such as tanks of liquids).  A simple metal or otherwise-minimally-protected back shell would work. 

At Mars,  the entry heat protection is far easier,  just because the velocities are about a third of,  to at most half of,  what we have to deal with at Earth.  For example,  even for direct entry from interplanetary trajectories,  the entry interface velocity is only around 5.6 km/s.  Low density ceramics are thus quite feasible,  even at stagnation conditions,  and even at high ballistic coefficients. 

Entry trajectories must be shallow at Mars,  just because the air is so thin.  If you come steep,  you are still way hypersonic when you smack the surface,  even in the lowlands.  This more-or-less rules out large delta-vee burns for de-orbit purposes.  A big deorbit delta-vee is invariably associated with a steeper entry trajectory angle,  that’s just the physics of orbital mechanics. 

Now, if you have a structurally-tough,  low-density ceramic,  then you have a reusable heat shield for a reusable “landing boat” at Mars.  That’s where my oddball experimental material has a huge amount of potential.  More so there at Mars,  than here at Earth. 

For entry from LEO,  one can use a winged or lifting-body shape instead of a blunt capsule.  It still needs a nose and leading edges that are as blunt as one can make them,  and these will require ablative protection.  Well away from stagnation zones,  the low-density ceramic refractories become feasible.  The more broadside you can fly during entry,  the more effectively-blunt you become,  and the lower the peak heating you must deal with. 

But (and this is a very,  very big “but”!!!),  angle of attack is severely limited by the structural problem of airloads ripping the wings off (or simple crushing breakup).  For shuttle this was 20-30 degrees.  Stray outside that AOA range (or almost any off-angle yaw or roll),  and you die.  One crew did die,  because of a hole in a wing leading edge.   

A final thought about LEO.  The retro deceleration burn from LEO (or LMO) is a far lower-precision thing than trajectory-adjusting burns during transit.  You don’t need precision-controllable liquid propellant thruster rockets for that.  The cheapest and simplest solution (and often the lightest-weight in a one-shot situation) is a small solid rocket motor. 

The retros on Mercury were solids,  and they worked quite well.  You get to design like the JATO bottles they used launching overweight airplanes,  for an application like that.  Which makes a difference.  They become simple “wooden rounds”,  as far as handling and logistics are concerned.  Actually,  very cheap. 

Anniversary of First Moon Landing

Today is the 44th anniversary of the first manned landing on the moon.  Neil Armstrong and Buzz Aldrin rode the lander to the surface,  and Michael Collins manned the spacecraft in lunar orbit.

This event had the same historical significance as Columbus's landing in the Caribbean half a millennium earlier.  Unlike then,  we have not capitalized on the moon landing.  Not in all the intervening 44 years has a human set foot on another world,  not even a return to the moon.

I suggest we make a holiday out of this date,  to celebrate the first moon landing the same way we celebrate Columbus Day.  It would be a fitting memorial to Armstrong and the rest,  if we did this. 

It might also help raise public awareness of the excitement and adventure of manned exploratory spaceflight.  The moon,  Mars,  the asteroids,  and uncountable places beyond,  all beckon. 

On the Asiana 214 Crash

There was something deadly wrong in the one cell phone video of the actual SFO crash that has surfaced.  With the electronic glide slope system unavailable (as it was),  visual approach is necessarily more hand-flown using the old-time glide slope lights.  Hand flying experience is critical,  and too many of these "bus drivers" don't get it. 

That plane was about a span too low,  and way,  way,  way too slow (near-stall nose-up attitude painfully obvious) at about 1/4 mile from the marks,  a furlong from the seawall.  Two somebodies in the cockpit clearly weren't watching the airspeed indicator. 

The low altitude can be dealt-with as long as you have adequate speed.  But low speed is inevitably an accident about to happen.

And it did.

GW


some second thoughts 7-10-13:

As I said,  nobody looked at the airspeed indicator.  Needle should have been hitting the flaps/gear white line,  or the stick shaker wouldn't have kicked in.  You'd think that with 3 or 4 pairs of eyes in the cockpit,  somebody would have looked at the IAS.  Apparently not,  and that's an artifact of over-automation in the cockpit.  Bus drivers,  not pilots.

 

I know nothing of the automatic electronics,  not even the damn radio,  but I do know stick-and-rudder flying.  I could have gotten the damn thing down in better shape than they did,  and I have not flown anything for almost 15 years now.  My weak spot would have been the flare:  not having any experience at that cockpit height above ground,  no feel for exactly where all the bits of the plane are located,  relative to me in that seat.

 

I've never held a pilot's license,  but stick-and-rudder flying does come easily to me  precisely because I was originally educated as an aeronautical engineer.  I know exactly how planes work.  Under the eye of an appropriate pilot,  in past decades I have flown two light plane types,  and two multi-engine types.  It wasn't hard. 

 

On the Train Wreck in Quebec

News reports on this are still quite confusing,  but here is what I have been able to determine as pretty much the facts:

The train was parked and "secured" near Nantes (uphill and 7 miles away from Lac-Megantic),  and the driver went to a hotel for the night.  In this instance,  "secured" seems to mean a running diesel engine powering the air brake system,  I'd guess by the two-pipe system,  such that the air brakes were set in the locomotive and all train cars. 

Then there was some sort of fire in the parked locomotive,  to which the Nantes fire department and a railroad engineering division person responded.  They put out that fire,  which seems to have been fairly minor,  and shut down the running engine.  They had a procedure to follow,  and they followed it.  Everybody went home.

About an hour later,  the entire train rolled away,  picking up speed gravitationally,  and entering Lac Megantic fast enough to derail and cause the disaster. 

This is my suspicion,  based on what I have read about train air brakes: 

When they shut down the locomotive,  they killed the air compressor.  Train car second pipe pressures then fell due to air leakage,  which is inevitable.  Reduced train line pressures would act to apply car brakes,  but not with leak-drained reservoirs!  The second line that keeps the reservoirs filled would have been off,  too,  with the locomotive shut down. 

The locomotive air brake would also have been disabled by bleed-down with the compressor off.  I'm not sure about sequence,  but sooner or later,  all of the air brakes would have failed from bleed-down.  Apparently this took about an hour. 

Apparently,  nobody thought to set at least one car's handbrake,  and this must not have been in the locomotive fire procedure that they were following in Nantes.  One or two cars' handbrakes could have secured the train,  even without any compressed air at all in the air brake system.  That could have prevented this disaster.

As for the tank cars themselves,  these were unpressurized-liquid cars,  and reported to be DOT-111 designs,  which are well-known to be thin-skinned and easily punctured.  They were hauling a light crude that comes from fracking operations in North Dakota.  It would have flammability characteristics closer to diesel than heavy fuel oil. 

The problem with closed flammable liquid tanks is overpressurization explosions when exposed to fire.  Those are not detonations like high explosives,  but they are still extremely violent.  Once the cars are thrown together in a wreck and some torn open,  the fire starts.  Explosions become inevitable. 

The disaster is better and easier to prevent,  than to fight after-the-fact.

Recommendation:  set some handbrakes if you shut off the locomotive on a parked train.  Leave a note to that effect in the cab before you go. 

GW

Second thoughts (later,  same day):

Being a freight,  this was most likely a one-pipe air brake system,  not a two-pipe.  It doesn’t matter to the outcome.  Turn off the air compressor (and they did),  and after a while,  every brake component bleeds to zero.  Once that happens,  brakes release. 

When the last brake releases (car or engine,  doesn’t matter),  the train is free to roll,  unless some mechanical handbrakes somewhere were set.  Once free to roll,  if parked on any sensible slope at all,  the train will roll away under gravity.  That is inevitable. 

One pipe system or two,  same outcome as already described.

Same preventative as already described:  set a handbrake or two,  more of them on a steep slope.  There’s plenty of time to release them while the cars’ air brake components “charge up”,  with the (running) locomotive brake set to hold the train. 

 Update 7-19-13:

It is now my understanding that the engineer was supposed to have set 10 or 11 of the handbrakes on this train.  Obviously,  this didn't work. 

There are then only 3 possibilities:  (1) he did not set them,  (2) he did set them,  but they were ineffective (for a reason of considerable interest),  or (3) somebody else released the handbrakes. 

I think the authorities have their work cut out for them,  finding out which of these three possibilities caused the disaster.  It is easy enough to blame the engineer,  but there are two other possibilities that must be eliminated before that is a credible action for anyone to take.

Neither of the other two possibilities is a very comfortable thought.  Yet,  they MUST be dealt with.  If they are not,  then any "final report" on this disaster is neither credible nor useful.