Saturday, August 31, 2013

Reusable Chemical Mars Landing Boats Are Feasible

I was initially quite skeptical that a single-stage chemical lander could be feasible for one-shot use, much less reuse-for-multiple trips,  with refueling.  But,  after some investigation with bounding analyses,  I have changed my mind.  These things are very feasible.  So,  we need not incur the difficulties of solid core nuclear propulsion,  in order to have a very practical ferry capability between surface and orbit at Mars.

Propellant Options
I do not know what propellants might really turn out to be manufacturable on Mars.  Right now,  I doubt anyone else really does,  either.  So,  I picked four combinations that might be “typical” of things others are considering.  The exact species were simply what was available in some old references on my bookshelf.  The four,  with rationale,  are:

Liquid oxygen – rocket-grade kerosene,  LOX-RP1,  well-developed technology
Nitrogen tetroxide – unsymmetrical dimethyl hydrazine,  NTO-UDMH,  well-developed, fully storable

Liquid oxygen – liquid methane,  LOX-CH4,  new,  easily-manufacturable?
Liquid oxygen – liquid hydrogen,  LOX-LH2,  highest-performing,  from ice?

Rocket Ballistics and Design

The data I had on-hand in my library allowed me to determine an estimate of rocket characteristic velocity (c*) as a power function of chamber pressure for all four selections.  I was able to scale the c* to an easily-achieved design goal of 500 psia (3.447 Mbar) for a medium-to-small rocket engine.  Expanding from that,  to 6 mbar backpressure,  gave me expansion ratio and ideal thrust coefficient,  read off a standard chart.  I corrected that by an appropriate nozzle efficiency reflecting realistic half-angles,  for single-engine uninstalled engine performance and gross geometry, derivable from a thrust specification. 
Once sized by a thrust requirement,  the throat diameter is known,  which with the expansion ratio and half angle sizes-out the bell dimensions.  I picked an arbitrary 10:1 area contraction ratio chamber-to-throat,  and sized chamber length off some empirical L* values.  Thus “reasonable” engine dimensions can be estimated,  once a thrust requirement is set.  The point is “getting into the ballpark”.

The Payload
I did my study around the concept of a fixed payload to be landed.  For an early mission,  this might typically be 3 men with a month’s supplies.  I included a guessed allowance for a rover car with a drill rig on it,  and allowances for surface exploration and experimentation gear,  plus a small inflatable “pup tent” in case they were too far from the lander to return.  I got 3.191 metric tons to cover this,  but that’s just a guess.  So also is 60+ cubic meters to contain it all,  plus some “living space” inside the lander. 

Flight Paths
These vehicles will have ballistic coefficients far larger than any of the lander vehicles so far delivered to Mars.  Investigations I have run came to the same conclusions as others:  such vehicles can aerobrake successfully,  but will come out of hypersonics at too low an altitude for chutes or ballutes to deploy,  much less do any good,  in the thin “air” of Mars.  Therefore,  hypersonic/supersonic retro-propulsion is going to be mandatory.  The vehicle simple rocket-brakes directly to touchdown,  once the entry hypersonics are over,  or at least mostly-over.  That kind of descent is illustrated in Figure 1 (below). 

These estimates are based upon a surface-grazing transfer ellipse from a 200 km altitude orbit (low Mars orbit,  LMO).  The de-orbit burn requires only a 50 m/s “delta-vee”,  something that attitude thrusters can provide.  LMO speed is 3.455 km/s at 200 km.  Speed at the 140 km interface is 3.646 km/s.  Speed at the local Mach 3 point varies some,  but falls in the 0.7 km/s range.  Allowing for heavy rocket-braking inefficiencies plus final near-hover for touchdown,  a good guess for the effective ideal “delta-vee” for descent is 0.9 km/s,  after getting some 2.9 km/s deceleration from aerobraking. 
The ascent is easier to rough-estimate.  The total velocity to be achieved is LMO speed,  with empirical “kitties” added to cover gravity losses and drag losses.  For rough-estimating clean rockets here on Earth,  5% of target velocity is adequate for each of both kitties.  The losses are less at Mars because of the weaker gravity and thinner “air”.  I simply ratioed-down the 5% figures by the surface ratios of gravity and density,  to a combined-loss “kitty” of 1.94%.  The ascent ideal “delta-vee” figure is then at most about 3.6 km/s,  as given in Figure 2 below. 

Those “delta-vees” add (for no reduction in payload upon ascent) to a total two-way “delta-vee” capability of about 4.5 km/sec.  For estimated installed specific impulse performances,  excepting hydrogen,  the resulting propellant fractions then fall near 75%. 
The Vehicles

Each propellant combination leads to a different vehicle size-out.   My concepts are based around a guessed constant 20% inert weight,  to cover both structural and landing equipment items,  but hopefully with enough robustness to provide a significant service life in re-use.  Once a propellant fraction has been determined,  it and the inert are deducted from one,  determining the payload fraction.  The known payload weight then sets the entire weight statement.  The key is thus determining the propellant fraction,  from performance parameters and the total “delta-vee” requirement. 
My vehicle layout concept is based on historical US capsule shapes:  blunt heat shield with a more-or-less conical afterbody.  Heat shield shape is spherical-segment,  with a radius of curvature equal to the diameter.  I used a conical afterbody in the 20-to-30 degree half-angle range,  with a cylindrical extension,  since cylindrical tanks and pressure vessels are easier to build than conical ones.  The flight control station is at the end of the cylindrical section,  for really good visibility. 

The structure would be a deck frame with attached heat shield,  and four extendible landing legs.  Height-to-stance width is near one,  for good stability.  In the center is a cylindrical sealed compartment containing 4 canted engines (I chose 10 degrees arbitrarily) to “enforce” retro plume stability during descent.  With a sealed engine compartment,  there can be no throughflow through wide-open ports in the heat shield for the engines.  That eliminates the need for port covers,  and for swapping ends during hypersonic flight. 
A small makeup-massflow/coolant-flow might be needed to balance the volume-filling transient as the vehicle descends,  in order to prevent intrusion of any entry plasma through the open ports.  That was not analyzed here.  But,  it should be noted that this method of preventing throughflow (with a sealed engine compartment) should work equally well,  whether the engines are firing,  or not!

I arbitrarily picked an installed thrust sizing such that the vehicle accelerates at two standard gees,  at its maximum ignition weight.  I also assumed 4 engines in the cluster.  Correcting for cant angle and four engines,  produces the individual engine size-out,  and thus the necessary engine compartment dimensions,  and all the performance parameters. 
The required quantity of propellants and their densities sizes the tankage volume to stack on top of the engine compartment.  The 45-degree cone,  on top of all of that stack,  is the crew flight station.  The conical segment around the engine compartment is the crew living space plus cargo volume plus compartmentalization for airlock purposes:  60+ cubic meters.  I assumed that some of the conical shell panels are hinged at the deck line,  so that they could also be used as unload ramps.  Only the LOX-LH2 vehicle sized-out such that the conical segment had to extend partway up the cylindrical propellant tankage section,  a consequence of its lower mass (but similar overall volume) relative to the other three.  That vehicle had the smallest engine compartment,  by far. 

Figure 3 below shows rough vehicle layout and dimensions,  all four being roughly the same overall shapes to within a fraction of a meter.  The weight statements are quite different,  as a function of propellant selection,  as shown in Figure 4.  Note than none of these vehicles could ride assembled to Earth orbit inside existing payload shrouds,  which are around 5 m diameter max. 
These things will have to be assembled on-orbit in low Earth orbit (LEO) from docked and assembled components,  and then sent to Mars.  I would suggest each lander push its at-Mars propellant supply to Mars as an unmanned cluster vehicle,  via min-energy trajectory.  The quantity of propellants each lander pushes to Mars then depends upon the scope of planned on-orbit-based operations.  Determining that is out of scope here.

Rough Performance Estimates
These vehicles could be operated as orbit ferries in either of two ways:  (1) on-orbit basing,  meaning refueling from supplies in LMO,  and (2) surface-basing,  meaning refueling from supplies manufactured on the surface.  In the first case,  entry is at very nearly maximum mass,  followed by ascent at reduced mass.  In the second case,  ascent at maximum mass is followed by entry at much-reduced mass.  These two cases very much affect the entry ballistics,  resulting in much-different final rocket-braking needs. 

On-orbit basing is the more stringent descent case.  My analyses indicated rocket-braking requirements in the 3-to-5 gee range,  if braking was delayed to the local Mach 3 point.  This violated considerably the two-gee engine sizing assumption.  Indicated “hover time” at 200 m was barely adequate at 55 sec to cover the touchdown.  This outcome simply indicates that rocket braking must start earlier,  closer to the entry max deceleration gees point,  which is at substantially-hypersonic speeds.  There is no reason this could not be done,  once the hypersonic/supersonic retro-propulsion approach is adopted at all.  I did not re-analyze this change in detail,  having already established the basic feasibility. 
With surface-basing,  the descent rocket braking requirements (waiting to local Mach 3) all fell within the two-gee intended design.  The 200 m hover times were at least twice those of the other case,  for all four propellant options. 

For both cases and all four propellant options,  entry peak deceleration gees fell in the 0.71-0.73 range.  It is the 2-gee rocket braking that is the highest gees the crew must withstand during descent.  Accelerations upon ascent were not analyzed,  but should be comparable to 2 gees,  assuming throttleable engines.  Clearly,  the crew should arrive at Mars fully physically-fit,  and able to endure a few minutes at 2 gees,  seated,  with full human functionality. 
This implies that their transit vehicle should provide artificial gravity (by spin) at 1 full (Earth-normal) gee.  That is an issue for manned transit vehicle design,  out of scope here,  but very definitely needing attention called to it.   

Employment of the Landing Boats
These craft could be employed flying multiple missions,  either on-orbit based,  or surface-based.  It is easy to imagine a first manned exploration mission to Mars,  where the landing boats are operated on-orbit,  using propellant supplies brought from Earth.  At the end of the mission,  the landers and any remaining propellants would be left in Mars orbit “for subsequent missions to use”. 

A subsequent mission with the objective of establishing a more-or-less permanent base or outpost,  might use them differently.  The initial landings,  to ferry vehicles and equipment down,  might use on-orbit propellant,  with subsequent flights using propellants manufactured on the surface.  Once that capability exists,  and we are interested in only one (or maybe two) specific outpost sites,  the landers would be operated in surface-based mode,  probably for the remainder of their service lives. 
A variation on that scenario would be to accomplish both objectives in one manned mission to Mars,  something feasible because the stay time at Mars until the orbits are “right” for the return is over a year long.  Initial on-orbit based explorations are done,  until a site is identified where large quantities of propellants can be manufactured fairly rapidly.  Then the vehicles and equipment are all transferred to that site,  for surface-based operation,  from what will become a more-or-less permanent outpost.  This presupposes that propellant can be manufactured in 20-ton lots on a timescale of a few weeks,  on that first manned trip.  If that is not true,  we are inevitably reduced to the first scenario. 

I would suggest that we plan for the two-objectives-in-one-mission scenario,  but with sufficient on-orbit propellants to fully support the fallback position. 
About the Heat Shield

The entry analyses yielded a worst-case peak stagnation heating rate of roughly 5.5 W/  This is low enough to allow the use of black-surfaced low-density ceramic heat shield materials on the windward surfaces,  even at the stagnation point.  White-surfaced low-density ceramics can be used on all lateral and leeside surfaces. 
However,  because these vehicles land on natural regolith surfaces,  there is the dead certainty of suffering dirt and stone impacts to the heat shield,  due to rocket back-blast effects at takeoff and touchdown.  Low density ceramics far less fragile (and far less labor-intensive to maintain) than the well-known Shuttle tile are thus demanded. 

There might be one:  my oddball experimental ceramic-ceramic composite appears to have the necessary toughness,  with no “show-stoppers” anticipated to completing its development.  This material is described in the posting dated 3-18-13 and titled “Low-Density Non-Ablative Ceramic Heat Shields”.  My recent well-received paper at the 16th Annual International Mars Society convention (in Boulder,  CO) also covered this same proposed material.   
Making the Propellant Selection

That topic is beyond scope here.  I do not know what propellant combinations might actually prove practical to manufacture on Mars.  I doubt anyone else really knows yet,  although many might claim to know.  That answer needs to be found first,  so that the landers we design and send,  on that first manned mission,  are compatible with what we can actually produce there.  That way,  they can serve for quite a while.  Making that selection is the fundamental pacing item for picking a lander design approach,  and then making it a flight-ready vehicle. 

1.       The first priority is to decide “for sure” what propellant selection could “best” be manufactured on Mars.

2.       Next,  the “landing boat” design given herein,  matching that propellant selection,  should serve as the design start-point for an actual “landing boat” design.  This design should be built,  tested,  developed,  and readied for operational use. 

3.       The manned Mars mission,  or sequence of missions,  should employ this “landing boat” in the relevant role (or roles),  and these craft should be left there for future uses. 

4.       This “landing boat” should be designed with the maximum possible expected service lifetime,  far beyond the needs of one mission,  so that it may serve subsequent missions or roles with local refueling,  for as long as possible. 

5.       There is no reason this “landing boat” development could not be started right now,  and thoroughly tested in low Earth orbit,  just as the lunar lander was.  The pacing item is (again) deciding which propellant selection could actually be manufactured upon Mars. 

Figure 1 – Descent Trajectory Assumptions and Requirements

Figure 2 – Ascent Trajectory Assumptions and Requirements

Figure 3 – Rough Overall Vehicle Dimensions,  Without Internal Layout Details

Figure 4 – Weight Statements and Performance Parameters by Propellant Selection

Update 9-9-13:  Some Further Thoughts About Selecting the Right Propellants

I really don't think kerosene is something we could practically manufacture on Mars,  but it might be somewhat representative of a hydrocarbon heavier than methane,  that we might dream up a process for.  It is a very well-known technology.  But,  I'd bet we can find a way to ignite or keep-unfrozen any of these choices,  though. 
I really don't think NTO or any of the hydrazines might actually be practically manufactured on Mars,  without a source of fixed nitrogen.  That's a huge obstacle there,  as far as I know.  But,  we already know those propellants can be easily stored,  and we have had engines that re-light multiple times in vacuum with them,  for decades now.  That's pretty much the technology of the shuttle OMS maneuver engine pods. 

I suspect LOX-LH2 would actually be the "easiest" to manufacture on Mars,  using mined ice and electrolysis as the basis.  LOX is not too much trouble to liquefy and store;  LH2 is much trickier to do,  with the ortho vs para form problem perhaps now the easiest problem of several to resolve. 

On Mars,  the truly fundamental problem is "where is the ice deposit big enough to be worth mining?"  We now know for sure that Mars has lots of water still (in the scientific sense),  the trouble is that it's just not "everywhere".  The kind of ice lenses Phoenix found near the pole is not the kind of deposit that supports practical mining and manufacture.  What we need is a buried glacier 10+ meters thick and many,  many km in lateral extent. 

BTW,  it'll be subliming as we dig it out.  Every mine hole has to be regolith-buried when not in use.  There will be one whale of a lot of regolith-moving operations involved in this activity.  The machines will look like heavy mining and road-building equipment.  That takes a big lander,  even if shipped in small pieces and assembled on site.  These things will not be carried by a series of Apollo-like dinky-little landing modules.  No way.  We need real "landing boats" of very significant size. 
They're not gonna fit existing payload shrouds for launch to LEO for this mission.  Something else to think about. 

We have orbital observations of where some such buried glaciers might be (emphasis on "might"),  but we have absolutely no ground truth about it.  I have never seen a robot probe design capable of determining that kind of ground truth,  either.  So,  if we are going to plan on making LOX-LH2 to return,  where do we land? 

Tough question.  We have to be close enough to walk to the ice,  or it ain't gonna work.  We're talking front end loaders,  bulldozers,  and large pressure-vessel process machinery here,  with maybe even some pick-and-shovel work by more than 2 men.  Long range transport is simply out of the question,  that first time up with propellant manufacture. 
As for making methane out of water,  and the CO2 in the "air",  the low inlet densities make all your machinery (whatever it is) look very large and heavy and energy-intensive,  compared to what we are used to here at home,  by about a factor of 14.  My guess is you can make 1's,  not 100's,  of kg per day.  You'll not accumulate enough to return a crew (tens of tons),  not even in a year's stay,  even if it doesn't break down or encounter unexpected problems. 

And you will encounter unexpected problems (lots of emphasis on "will").  Done robotically before the men arrive offers a potential way out,  except that robots-as-we-know-them-today are simply inept at solving unexpected problems.  Put the men there to solve those problems,  and you are right back to the inability to accumulate tons of propellants in time.  Plus,  with LOX-CH4 you still have to solve basically the same water problem as LOX-LH2,  to get the oxygen and the hydrogen. 
So I dunno which one to try.  And I don't yet see much of a path to resolving this in time for a mission in the 2030's,  much less the 2020's we'd all like to see.  NASA has no plans to send the right kind of probes that could locate the propellant-making resources.  I don't see anybody else sending the right kind of probes,  either.

That puts me back to the costly-but-sure-thing concept:  first mission relies on propellants-sent-from-Earth.  Which means it is an LMO-based mission,  sending down multiple ferries to multiple interesting sites,  and emplacing the machinery to experiment with propellant manufacture at the most promising ones after the men return home.  Leave the ferries in a higher Mars orbit,  with whatever propellant is left over,  for the next mission to use.  What's the point of going all that way with men,  and only making one landing?  That's really dumb!
Meanwhile,  we have to guess which propellants might actually be made on Mars most practically,  and build the first-mission ferries to use that.  That way subsequent missions (including planted bases) can refuel and re-use the same ferries.  Right now,  I'd guess LOX-LH2 from ice.  But with an engine compartment big enough to accommodate being refitted with different engines.  And with compartmentalized tankage to accommodate being re-plumbed for different propellants.  That's heavier,  and so is structural robustness necessary for long-term reusability.  My assumptions of inert structural weight 20% are quite likely too low. 

 Update 9-11-13:

I have had some conversations with John Strickland about his designs versus mine.  Coming from very different starting points,  his results for lander vehicles and mine are amazingly similar.  This study of mine is a lot more realistic than should ever be expected of a bounding calculation.  (And that's what it is,  so don't read too much into all the nitty gritty little details).   

Update 9-22-13:

I'm beginning to think,  since all 4 of my vehicle rough-outs turned out to be about the same overall physical size,  that a "good" design might be one based around LOX-LH2,  but with enough space internally to be reconfigured and re-engined for LOX-CH4.  Comments?  Ideas?  Please weigh-in!

My best guess is that this LOX-CH4 combination is the "most likely" in-situ-produced propellant combination,  long term.  Short term,  I really think it might be LOX-LH2 from mined ice,  with nothing but electrolysis and liquefaction.  That is the simplest and most direct combination we have.  It is only restricted by where significant ice is actually buried on Mars. 

Basically,  vehicle size is fundamentally "set" by the payload mass to be landed.  My concepts are for "well-empowered" explorers,  not permanent base builders.  Please weigh-in with payload mass ideas for the follow-on base-building missions.  I know a lot less about that. 

The real question to answer here is:  what are we really going to do with men on Mars?  Explore?  Build bases?  Both sequentially?  Both at once?  The answer makes a huge difference to the mission approach,  architecture,  and component designs.  Overwhelming,  actually. 



Wednesday, August 28, 2013

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. 

Tuesday, August 20, 2013

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. 


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.


Sunday, August 4, 2013

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/,  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.