A Question For Readers

Here is a question for my readers who are particularly interested in the space travel stuff that I do.  Would you be interested in participating with others of like mind,  working on these kinds of ideas?  If so,  you might visit,  or even join,  the community that is on the New Mars forums,  part of the Mars Society’s efforts. 

That link is http://newmars.com/forums/

I am a frequent contributor there,  and we on those forums are looking for more folks interested in trying to contribute toward any possible viable ideas.  Give it a try.  Let me know here what you think,  if you go visit there.  Please use the comment option here.  I only remove obvious spam. 

GW 

Pertinent Funny

This one speaks for itself,  but is pertinent to so many things I see going on.  Enjoy.  My wife found this for me somewhere on her Facebook.  

Refining Proposed Suit and Habitat Atmospheres

I came up with the design analysis of suit and habitat atmospheres posted in Ref. 1,  and then developed a simplified and organized spreadsheet model,  to implement that design analysis procedure,  all in one convenient place.  This model uses a long-term hypoxia criterion developed from data in Ref. 2 for the habitat,  and two short-term hypoxia criteria for the minimum-pressure suit,  from pilot oxygen mask requirements.  I developed a fire danger criterion for the habitat out of oxygen concentration,  per its use in Arrhenius-type reaction-rate models. The “no pre-breathe” criterion is NASA’s,  via the USN.

The fully-compliant habitat and min-pressure suit atmosphere values of Ref. 1 are now the default case in the spreadsheet model.  This was reported in the Addendum to Ref. 1.  I have since done two further analyses,  denoted “work case 1” and “work case 2”,  as their own worksheets in the spreadsheet.

The default case at 0.45 atm and 45% oxygen (by volume) in the habitat,  produced a recommendation well in excess of the long-term hypoxia criterion,  even leaked down to 0.40 atm (some 11.111% lower). It was compliant with the fire danger criterion,  and produced a 3.031 psia pure-oxygen suit proposal that is compliant with the fully-cognitive short-term hypoxia criterion,  even if leaked down by 10% on pressure.  This is the lowest suit pressure that meets no pre-breathe.  Anything higher also requires no pre-breathe.  This is a very good combination,  but I wanted to see if I could do even better.

For “work case 1”,  I reduced the design (max) habitat pressure to 0.40 atm,  and increased the oxygen to 50%,  with a 10% pressure leak-down specified for both suit and habitat.  This still meets the long-term hypoxia criterion for the habitat,  even leaked down,  and it still meets the fire danger criterion,  by a very slightly better margin.  But it produced a min pressure suit option that failed to meet the fully-cognitive short term hypoxia criterion entirely,  and also failed to meet the bare survival hypoxia criterion when leaked down. 

I had to separately raise that suit pressure back up to 3.013 psia pure oxygen,  before it met the fully-cognitive short-term hypoxia criterion,  even leaked down.  This suit also needs no pre-breathe.  Paired with the upgraded min suit pressure,  this is also a good combination,  although it wastes some of the pre-breathe margin.  I added a separate suit upgrade calculation off to the right,  in this worksheet.

So then I ran “work case 2”.  I started getting acceptable suit pressures at about 0.43 atm habitat pressure,  and I fully met the habitat long-term hypoxia and fire danger criteria,  at just about 43.5% oxygen.  I had to hunt around a bit on both habitat pressure and oxygen percentage,  before settling on these values.  They resulted in a min suit pressure that was just a bit lower than the default case or “work case 1” at 2.975 psia,  but it still met the short-term fully-cognitive hypoxia criterion,  even when leaked down 10%.  This is the best combination I have yet found.  The separate suit upgrade calculation is also in this worksheet,  but was not needed.

The default case is Fig. 1,  which is also Fig. 9 in Ref. 1,  “work case 1” is Fig. 2,  and this best-version-yet “work case 2” is Fig. 3.  The previously most recent posting (prior to Ref. 1) about this subject is Ref. 3.

Spreadsheet Availability and Function

If you want a copy of the spreadsheet file,  please contact me by email.  As it says in the user instructions on the worksheets I created,  I recommend that you keep these example cases unchanged as templates.  Copy one of them to a fresh worksheet and do your design analysis there. 

If you copy “work case 1” or “work case 2”,  you get the suit pressure rework calculations as well,  off to the right of the main design analysis.  That is only necessary if your min suit pressure falls in a range that violates the short-term hypoxia criteria.  I did not put the revised suit calculation block on the “default case” worksheet.

If you instead want to create your own calculations,  just remember this critical point:  to get wet in-lung oxygenation,  you must first subtract-off the water vapor partial pressure to get the total partial pressure of the breathing gas inside the wet lungs.  Only after that is done do you get to apply the breathing gas volume percentages to that total partial pressure of breathing gas in the lungs. 

My calculations start with a proposed habitat atmosphere at some dry total pressure,  with a volume percentage of oxygen in it,  and also the assumption that it is a two-gas mix of just oxygen and nitrogen.  That produces the dry breathing gas partial pressures of oxygen and nitrogen. 

I reduce that total pressure by the vapor pressure of water at human body temperature to find the partial pressures of the breathing gas in the wet lungs,  and apply the volume percentages to that reduced value,  to get the partial pressures of oxygen and nitrogen in the wet lungs.  The partial pressure of oxygen in the wet lungs compares to the long-term hypoxia criterion of min 0.14 atm.

I do a molecular weight calculation to determine the mass fraction of oxygen in the mix,  which multiplies the dry breathing gas density to produce the oxygen concentration as mass per unit volume,  for comparison to the fire danger criterion of max 0.275 kg/m³,  for warm dry sea level air at 77 F = 25 C.

The partial pressure of nitrogen in the dry habitat atmosphere gets divided by the NASA/USN “no pre-breathe” factor of 1.2,  to produce the minimum pure oxygen suit pressure you can use,  and still avoid a pre-breathe time requirement.  This gets the vapor pressure of water subtracted to find the wet in-lung partial pressure of oxygen.  That gets compared to the short-term hypoxia factors:  min 0.12 atm for full cognitive capability,  and min 0.10 atm for bare survival.  (Somewhere under about 0.08 atm is the “certain death-by-hypoxia” point,  although such exposure does take significant time to injure or kill.)

It is entirely acceptable to find a habitat atmosphere at somewhat lower pressure and slightly higher oxygen than my best recommendation (“work case 2”),  that meets long-term hypoxia and fire danger criteria,  yet the resulting minimum pure oxygen suit pressure fails to meet the short-term hypoxia criteria (that is exactly that happened in my “work case 1”). 

That minimum suit pressure is just a lower bound on what you can design your suits to have.  You can always design your suits to a higher pressure than this lower bound,  to meet the hypoxia criteria.  They will always then satisfy the “no pre-breathe” criterion.  That is exactly what I did in “work case 1”,  and it is precisely why I added the suit pressure redesign block out to the right of the main calculation block.

References

#1. G. W. Johnson,  “Habitat Atmospheres and Long-Term Health”,  posted 1-1-2022 to http://exrocketman.blogspot.com

#2. Martin Enserink,  “Hypoxia City”,  a science news article published in the journal magazine “Science”,  volume 365,  Issue 6458,  dated 13 September 2019,  as published by the American Association for the Advancement of Science (AAAS).   

#3. G. W. Johnson,  “Suit and Habitat Atmospheres 2018”,  posted 16 March 2018 to http://exrocketman.blogspot.com 

Figure 1 – Default Case is Best Case From Ref. 1 (0.45 atm at 45% O2)

Figure 2 – This Is the “Work Case 1” Worksheet,  Now In the Spreadsheet File (0.40 atm at 50% O2)

Figure 3 – This Is the “Work Case 2” Worksheet,  In the Spreadsheet,  and The Best Yet (0.43 atm 43.5%)

Addendum:  “Rule of 43” for Habitat and Suit Atmospheres

Here’s a design combination that is really easy to remember,  and yet gets just about as good an answer as the fully optimized form.  The optimum case had a habitat atmosphere that was 43.5% oxygen at 0.43 atm pressure.  It produced a minimum oxygen suit pressure of 2.975 psia.  The habitat satisfied the fire danger criterion,  and the long-term hypoxia criterion,  even leaked down 10%.  The suit met the no pre-breathe time requirement,  and the fully-cognitive short-term hypoxia criterion,  even when leaked down 10%.  It would be more easy-to-wear as a gas balloon design than current NASA suits,  by far!  It would be even more feasible and easy-to-build as an MCP suit than what Dr. Webb did in the 1960’s. 

The “rule of 43” case gets very similar results,  but is far easier to remember.  It uses a habitat atmosphere that is 43% oxygen at 0.43 atm pressure (both “43”).  It meets the fire danger criterion,  and meets the long-term hypoxia criterion if leaked down no more than 9.5% (it just barely fails at 10%).  The min suit pressure for no pre-breathe time comes out just a tad higher at 3.002 psia pure oxygen,  and meets the short-term fully-cognitive hypoxia criterion at 10% leaked-down.  Like the optimum case,  this would be far easier to wear as a as balloon suit,  and far easier to build as an MCP suit.

Figure 4 is the “rule of 43” combination,  and Figure 3 above is the optimum combination that I found earlier.  These were done with the spreadsheet tool I developed,  and in just a matter of less than an hour,  iterating through several possibilities where the atm of pressure and the oxygen percentage were the same numbers.

Figure 4 --  “Rule-Of-43” Design Case At 0.43 Atm Pressure And 43% Oxygen

These two cases are so close,  that I see very little difference between them.  If the objective of “something easy to remember” is as important as I have been told it is,  then this “rule of 43” design is the one you really want.  Its no pre-breathe min suit pressure is very slightly higher,  and its habitat pressure leak-down percentage isn’t quite the full 10%,  but that doesn’t really matter.  Both are in the very same ballpark,  with the differences out in the decimal places. 

The main point here is to get into that ballpark,  so as to reduce the min suit pressure for no pre-breathe way below NASA practice,  so that easier-to-wear gas balloon suits become feasible,  and that even easier-to-build MCP suits become possible.  These suit pressures are quite adequate,  but are far below what NASA and its favored contractors have been using (3 psia vs over-4.2 psia). 

You find out how adequate these lower suit pressures really are,  once you generalize the health and oxygen mask altitude criteria to wet in-lung oxygen partial pressures.  You need that generalization of those criteria,  in order to extend them correctly to lower pressures and higher oxygen percentages,  than those of Earthly air.  You also need a fire danger criterion cast in the mass/volume chemical concentration format.  And,  you need suit short-term hypoxia criteria based on Earthly use of oxygen masks for pilots at high altitudes.

Utter-Minimum Suit Pure Oxygen Pressures

I used the “work case 2” suit upgrade calculation block to investigate just how low a suit pressure was safe,  using the short-term hypoxia criteria.  Remember,  a wet in-lung oxygen partial pressure of 0.12 atm supports a fully-cognitive wearer.  0.10 atm supports survival without full cognition:  the wearer may well be somewhat nonfunctional mentally. 

Figure 5 is what I get if I require the fully-cognitive hypoxia criterion to the suit in the 10% leaked-down state.   Figure 6 is what I get if I only require the fully-cognitive hypoxia criterion to the design pressure;  leaked down 10%,  it fails fully cognitive,  but still satisfies bare survival.  The lesson here is that suit pressures as low as 2.675 psia will be quite adequate for fully-cognitive wearers.  2.407 psia will save life,  even if the wearer is mentally not fully functional.

Figure 5 – Min Suit For Fully-Cognitive When Leaked-Down 10%

Figure 6 – Min Suit For Fully-Cognitive Only At Design Pressure

A word of caution:  these utter-minimum pressure suit designs cannot be used indiscriminately with the two long-term habitat atmospheres identified so far (0.43 atm and 43.5% O2,  and the “rule of 43” design with 0.43 atm and 43% O2).  The utter-minimum pressure designs violate the min suit pressure specs for no pre-breathe time,  because the ratio of habitat nitrogen partial pressure to suit design pressure exceeds the 1.200 criterion. 

I include these utter-minimum suit design specs here,  to show what is actually feasible for adequate life support and mental functionality in pure oxygen suit designs,  when those designs are independent of a habitat pressure that must meet a long-term hypoxia criterion (for the safety of pregnant women and unborn/newborn children). 

Habitat Atmospheres and Long-Term Health

Update 1-11-2022:  revised Figures 1 and 2 to show correspondence of curve fit and criterion.

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

I had originally intended to do this as an update to an existing article (Ref. 1),  but since have changed my mind.  This is different enough to warrant being its own article. 

The data and recommendations in Ref. 1 are based on experiences for relatively short-term exposures,  a few days or weeks similar to Apollo and Skylab for habitat atmospheres,  and a few hours for low-pressure pure oxygen in space suits.  Those results are based on the wet in-lung inhaled oxygen partial pressure in the lung gases,  set by concentrations available to pilots flying aircraft,  and there’s nothing wrong with that approach for space suits!  But as it says in the article,  for very long-term exposures,  and especially for pregnancy and childbirth,  other constraints unknown to me then,  may apply. 

That is why in Ref. 1,  I recommended a general habitat atmosphere compatible with no pre-breathe time getting into an oxygen suit,  and also compatible with no greater fire danger than in warm sea level Earthly air.  But at the same time,  pregnant women and young children were limited to a portion of the habitat that had another atmosphere more similar in oxygen content to Earthly air,  at a higher pressure.

I have recently run across some very pertinent information regarding long term exposure risks,  and pregnancy and childbirth risks,  for life at high altitudes in Earthly air.  These would have a direct bearing on the atmospheric composition and pressure in any ship,  space habitat,  or any habitat upon some other celestial body.  I found this in an older issue of AAAS’s journal “Science” (Ref. 2).  It deals with health investigations made in the Andes,  up to altitudes of 5100 meters (16,700 feet).

What I do here is combine the Ref. 2 health vs altitude data with my wet,  in-lung oxygen partial pressure calculation approach from Ref. 1,  using a standard atmosphere table to relate altitude to total atmospheric pressure. The fire danger is based on the oxygen concentration,  expressed as mass per unit volume,  just as it was in Ref. 1.

Health vs Altitude Information From Ref. 2

The information in this reference is given in terms of altitude above sea level,  as it relates to hypoxia effects upon humans living at high altitudes on Earth.  Fundamentally,  oxygen diffusion across the lung tissues from the in-lung gases to the blood is driven by differences in oxygen partial pressures between the lung gases and the blood.  That is why I converted the relevant altitudes to oxygen partial pressures,  but only after letting water vapor displace some air inside the wet lungs (at human body temperature). 

There is also a carbon dioxide displacement effect,  from the carbon dioxide diffusing out of the blood into the lung spaces to be exhaled.  However,  this effect is far smaller than the water vapor displacement effect,  by at least an order of magnitude!  Thus it is safely ignorable for design purposes.  In contrast,  the water vapor displacement effect is significant at any pressure,  but very much more so at reduced breathing gas pressures,  precisely because the vapor pressure of the water depends on body temperature,  and just does not reduce with reducing total pressure.  It is simply a bigger fraction at lower total atmospheric pressures.

The “Science” article uses chronic mountain sickness (CMS) as the indicator for the relevant hypoxia effects.  The common external symptoms are dizziness,  headaches,  ringing ears,  sleep problems,  breathlessness,  palpitations,  fatigue,  and (particularly) cyanosis.  That cyanosis turns lips,  gums,  and hands a purplish blue,  easily observed by the most casual observer. 

Examined closer medically,  the hallmark of CMS was long thought to be a great proliferation of red blood cells,  to the point of making the blood more viscous.  This increase in viscosity raises the blood pressure,  especially in the arteries leading from the heart to the lungs.  It can also lead to enlargement of the left ventricle of the heart,  and thickening of its walls,  eventually causing heart failure.  

For short-term exposures (measured in days or less),  the treatment for hypoxia effects is supplemental oxygen,  followed quickly by a return to lower altitudes.  For long-term exposures (measured in months or years),  the only treatment is relocation to low altitude,  but some of the damage may be permanent!  According to the Ref. 2 article,  it has been known since the Spanish first tried to colonize the Andes in the 16^th century,  that the risks for eclampsia in pregnant women are elevated at high altitudes.  For the unborn and newborn at those same high altitudes,  the elevated risks are for low birth weight,  and for premature birth.  This problem of long term hypoxic exposure is really quite serious!

All that stuff has been known for a while,  but Ref. 2 also reveals some more recent information:  the increase in blood viscosity that precipitates CMS does not correlate with only the proliferation of red blood cells!  Everybody suffering hypoxia while living at high altitudes has more than the normal red cell count.  However,  only some of them show the CMS symptoms.  Exactly why this is true is not fully understood yet,  but there does seem to be some genetic components to it.  People living at high altitudes in the Himalayas,  the Andes,  and the mountains of Ethiopia all show different responses.

The Ref. 2 science news article indicates there are no CMS cases below 2500 meter altitudes,  which is 8200 feet.  8000 feet cabin altitude is now quite common in airliners,  when only a few years ago 10,000 feet was more common.  One thing to investigate would be the equivalent altitude for no CMS from the CMS case percentages versus altitude.  It should correlate to the 8200 feet criterion rather closely.

In South America (the focus of the Ref. 2 article),  a closer reading generates these data:  In La Paz,  the capital of Bolivia,  at 3600 meters = 11,800 feet,  some 6 to 8% of the large local population shows symptoms of CMS.  In the central Andean town of Cerro de Pasco,  at 4300 meters = 14,100 feet,  about 15% of those aged 30-39 showed CMS,  while some 33% of those aged 50-59 showed CMS.  (I simply misused those percentages as the upper and lower bounds for that population at that elevation.)  The article focused on the mining town of La Rinconada,  at 5100 meters = 16,700 feet!  This is a town of 50,000 to 70,000 people,  living and working under utterly deplorable conditions.  They show at least 25% with CMS,  and it may be a lot higher.  I arbitrarily used 50% as an upper bound.  It could be higher.

Figure 1 shows a plot of lower and upper bound percentages of CMS as a function of wet in-lung partial pressure of oxygen.  The “average” seems to strike the zero CMS level at about 0.13 atm for this population.  Per Figure 2,  this occurs at about 9000 feet.  Note that below 8200 feet (2500 m) the expected CMS rate is zero,  and the wet in-lung oxygen partial pressure is 0.14 atm (or higher).  The correspondence between these altitude criteria is very close indeed!  We can probably reliably use 0.14 atm wet in-lung oxygen partial pressure as a hypoxia criterion,  for either CMS or reproductive effects.

Figure 1 --  Occurrence of CMS vs Wet In-Lung Partial Pressure of Oxygen (Andes) UPDATED

Figure 2 – Wet In-Lung Partial Pressure of Oxygen vs Altitude UPDATED

For purposes of this article,  let us use 0.14 atm wet in-lung partial pressure of oxygen as the long-term limitation on a single habitat atmosphere composition and pressure for all residents,  including pregnant women and the unborn/newborn.  We will see what suit pressures and fire dangers result from this.  It pays to be conservative,  because of the long-term risks.

Habitat Atmosphere

To do this in the most straightforward way,  I made total habitat breathing gas pressure (in atmospheres) and habitat oxygen content (percent by volume) inputs.  This is for a two-gas mix of oxygen and nitrogen.  From these data,  using 0.06193 atm as the vapor pressure of water at human body temperature,  I computed partial pressures of oxygen and nitrogen in the dry breathing gas,  the displaced partial pressure of breathing gas in the (wet) lungs,  and the wet in-lung partial pressure of oxygen.  That procedure is intended to address the long-term hypoxia exposure issue,  for which we just determined the safe wet in-lung oxygen partial pressure needs to equal or exceed 0.14 atm. 

Running numbers to see trends,  I started at standard day sea level conditions,  with a two-gas “synthetic air” at 20.946 volume % oxygen.  I corrected air density at 1 atm to 77 F living temperature (which is where the 0.275 kg/m³ concentration came from).  Then I reduced the total synthetic air pressure by 0.1 atm increments down to 0.4 atm,  for a wet in-lung oxygen partial pressure about half the min value needed.  Then I raised the oxygen percentage to 40%,  getting very nearly the desired wet in-lung partial pressure desired.  0.4 atm at 40% oxygen is not quite adequate long term.

From that 0.4 atm - 40% oxygen point,  I then looked at 0.45 atm – 40% oxygen,  0.45 atm – 45% oxygen,  and 0.40 atm – 45% oxygen,  in order to “bound the problem”.  Note that this bounded problem is not unique:  one could use lower pressure and higher oxygen percentage,  and still meet the wet in-lung criterion.  Such choices might trip the fire danger limit,  though.

Fire Danger

I computed the molecular weight of the synthetic air as the sum of the volume percents of the gases times their molecular weights,  and then from this and the input volume percent oxygen,  the mass percent oxygen (as MWO2 * vol%O2 / MWsyn.air).  I used the habitat pressure and temperature to correct the sea level standard density of air in kg/m³,  and then multiplied that by the mass percent oxygen,  to compute the oxygen concentration in kg/m³.  That looks at the fire danger issue,  for which the oxygen concentration should be under the warm sea level air criterion of 0.275 kg/m³.

Min Suit Pressure

From the partial pressure of nitrogen in the habitat atmosphere,  reducing that by a factor of 1.2 is the “no pre-breathe” criterion for minimum pure-oxygen suit pressure.  At the higher habitat pressures nearer 1 atm with real synthetic air,  this produces ridiculously-large min suit pressures!  But as the habitat total pressure dropped near 0.4 atm,  and the oxygen percentage increased (driving down nitrogen even further),  the min suit pressures became quite attractive.  Higher suit pressures are compatible with no pre-breathe time,  but not suit pressures lower than this computed minimum.  I also computed wet in-lung oxygen partial pressure for that min suit pressure,  and for 10% leaked-down.

Results

I chose as the design value 0.45 atm total habitat pressure (which is 6.6132 psia) at 45% oxygen by volume (which is 2.9759 psia partial pressure of oxygen and 3.6373 psia partial pressure of nitrogen).  That is very similar to,  but not exactly the same,  as the habitat mix of Ref. 1.  This selection resulted in a wet in-lung oxygen partial pressure of 0.1746 atm,  which far exceeds the adopted 0.14 atm long-term exposure criterion.  Even with a leak-down to 0.40 atm,  the wet in-lung oxygen is still 0.1521 atm,  which significantly exceeds the criterion.  As long as long-term hypoxia is avoided this way,  there should be no troubles with CMS,  or with pregnancy and birth beyond the “normal occurrence” rates.

The computed oxygen concentration at 0.45 atm – 45% oxygen is 0.257 kg/m³,  which is under the 0.275 fire danger criterion.  Leaked down to 0.40 atm – 45% oxygen,  it is only 0.228 kg/m³.  The fire danger is similar to,  but very slightly less than,  that of sea level warm Earthly air. 

The minimum pure oxygen suit pressure for the 0.45 atm – 45% oxygen design point is quite attractive at 3.031 psia (157 mmHg),  although that is slightly higher than the suits explored in Ref. 1,  which met short-term hypoxia exposure criteria.  Any suit pressure at or above this value meets the no pre-breathe criterion for the selected design habitat atmosphere.  This is well below current NASA suit pressures,  and is still quite similar to the lowest altitude-equivalent suit recommended in Ref. 1 (2.945 psia = 152 mmHg).  That suit had a wet in-lung partial pressure of oxygen of 0.1385 atm,  almost meeting the long-term hypoxia criterion adopted here. 

In the event of suit leak-down,  there is still sufficient wet in-lung oxygen partial pressure to avoid short-term hypoxia effects,  as already explored in Ref. 1.  There is no pre-breathe or decompression time needed to go from an oxygen suit back to the habitat atmosphere,  even if leaked down 10%.   

See the data table in Figure 3,  for all the habitat atmosphere numbers run for this study. 

Figure 3 – Spreadsheet Image of Numbers Run for this Study

Recommended Suit Design Criteria to Match

The minimum-pressure altitude-equivalent suit design,  per the methods of Ref. 1,  is 7.72 kft (2350 m) altitude,  for which the pure oxygen suit pressure is 0.20625 atm = 3.031 psia = 157 mmHg.  The resulting wet in-lung oxygen partial pressure actually meets the long-term hypoxia criterion at 0.144 atm vs the 0.140 criterion.  If leaked down 10%,  the suit pressure is still 0.1856 atm = 2.728 psia = 141 mmHg.  That wet in-lung partial pressure of oxygen is 0.1237 atm,  which corresponds to Earthly air at 11.36 kft (3460 m).  This is still quite safe for short exposures,  as explored in Ref. 1.  See Figure 4.  

Figure 4 – Min Suit Pressure Data for No Pre-Breathe Time

Conclusions (see also Fig. 5,  6,  and 7):

#1. Use a long term habitat (or ship) atmosphere of 0.45 atm (6.61 psia) at 45% oxygen by volume,  in a simple two-gas composition with nitrogen.  This atmosphere is safe from a fire danger standpoint,  and more-than-meets the wet in-lung oxygen partial pressure criterion to avoid long term hypoxia effects,  like CMS.  Presumably,  this also avoids aggravating complications of pregnancy and childbirth beyond the normal occurrence rates.  It is still safe,  even if leaked-down to 0.40 atm at that same 45% oxygen.

#2. The minimum pure-oxygen suit pressure that does not incur a requirement for pre-breathe time with this habitat atmosphere,  is only 0.20625 atm = 3.031 psia.  This is a 7.72 kft altitude-equivalent design,  and is no worse than 11.36 kft equivalent when leaked-down 10%.  Any pure-oxygen suit design with that pressure or higher,  is free of the pre-breathe time requirement,  and meets the long-term hypoxia criterion.  It does not meet that criterion if leaked-down 10%,  although such exposures should be quite short,  thus doing no permanent harm.  

Figure 5 – Summary of Conclusions

Figure 6 – How the Habitat Atmosphere Fares Versus the Long-Term Hypoxia Criterion

Figure 7 – How the Min Suit Fares Versus the Short-Term Hypoxia Criteria

Post Scriptum

Where the short-term hypoxia criteria came from is more of a judgement call.  It is based on varying military and civilian oxygen mask requitements for pilots,  and on some high-altitude human populations.  Pilots require full cognitive abilities.  Somewhere between where the USAF and the FAA require supplemental oxygen regardless of circumstances,  is that criterion.  For our purposes,  that is around 0.12 atm wet in-lung partial pressure of oxygen,  and it also matches what obtains at the highest altitude for which a vented pure oxygen mask is effective for more than a single handful of minutes. 

For bare survival with cognitive impairment after several minutes to an hour or so,  the rare max cabin altitude of 15,000 feet will serve well enough.  That is about 0.10 atm wet in-lung partial pressure of oxygen.  You could go to 20,000 feet equivalent at 0.08 atm wet in-lung partial pressure of oxygen and still survive,  but most lowlanders will be unconscious at this condition.  There are some acclimatized and genetically-adapted herders in the Andes who spend parts of their days up there,  but even they return to lower altitudes at night.

This is shown in Figure 8,  along with the so-called Armstrong limit,  where the water vapor in the lungs displaces all of the air or oxygen you are trying to breathe.  That is the so-called “vacuum death point”.  Actually,  the real hypoxia death point is lower,  somewhere in the 50-60,000 foot range.  Some jet pilots have reached 50,000 feet with only an oxygen mask,  but they were only there for mere seconds.

Figure 8 – Data From Which the Short-Term Hypoxia Criteria Were Obtained

References

#1. G. W. Johnson,  “Suit and Habitat Atmospheres 2018”,  posted 16 March 2018 to http://exrocketman.blogspot.com

#2. Martin Enserink,  “Hypoxia City”,  a science news article published in the journal magazine “Science”,  volume 365,  Issue 6458,  dated 13 September 2019,  as published by the American Association for the Advancement of Science (AAAS).    

Addendum:

Since completing the write-up of these design sizing analyses,  I have created a well-organized spreadsheet that does these calculations quickly,  all in one place.  This allows iterative fine-tuning of the pressure and oxygen percentage selections for the habitat atmosphere,  while concurrently monitoring the long-term hypoxia criterion,  the fire danger criterion,  and the adequacy of the min suit pressure value in terms of short-term hypoxia criteria. 

One should note that the main difference between this min suit pressure sizing procedure,  and that used in Ref. 1,  is that the suit is determined directly from the nitrogen partial pressure in the habitat atmosphere.  At mixture,  the max value of habitat pressure produces the max value of nitrogen partial pressure,  and that determines the max value of minimum suit pressure required for no pre-breathe requirement. The Ref. 1 method used an independent altitude-equivalent way of setting suit pressure,  which in turn added another level of iteration to the habitat atmosphere design process.

Also slightly different from Ref. 1,  I have added a leak-down option for the habitat pressure.  For a short transient,  it is OK to violate the long-term hypoxia criterion in the leaked-down state.  But,  if you set the habitat design (max) pressure and leak-down percentage such that the long-term hypoxia criterion is satisfied,  even in the leaked-down state,  then you have relieved yourself of the need for fine pressure control.  You then have a range of long-term acceptable pressures,  all at the same oxygen-nitrogen mixture.

Figure 9 shows an image of the new spreadsheet design analysis tool.  User inputs are highlighted yellow.  There are only 4 of them:  habitat pressure,  habitat oxygen percentage by volume,  the percent habitat pressure leak-down,  and the percent suit pressure leak-down.  All the rest is automatic.   Significant items or outputs are highlighted blue or green.  The user instructions are on the worksheet.  Contact me by email if you want a copy.

It is recommended that the default design analysis worksheet be maintained as-is.  Simply copy and paste it into a fresh worksheet to do your own design analysis.  That way,  if you screw up a cell formula,  you have the original available to restore it.  The default input data correspond to this article’s design analysis.  The odd-looking habitat leak-down percentage is that which produces exactly the leaked-down habitat pressure used in that analysis.

Figure 9 – Image of Spreadsheet Design Analysis For Habitat and Suit Atmospheres

 

Christmas Decoration

My wife got this for me off her Facebook.  It is one of the better things I have seen.  Enjoy.

Update 12-16-2021:  This was just too good.  We made our own.  It's hanging in the entryway.

The Seal Failure in the SRB that Doomed Challenger

I have real difficulty with the fact that,  even after all these years,  it is still necessary to explain to people what really destroyed Space Shuttle Challenger and killed her crew,  back in January 1986.  It was really two very seriously-bad upper management decisions at NASA,  one long before the launch: 

(1) to insist on poorly-designing the O-ring seal joints with 3 interacting serious errors,  and

(2) to fly soaked-out colder than had ever been tested,  when everybody’s engineers did not want to.

Background

First,  you have to understand what really happens in federal government contracting.  There is only one customer,  and he thinks he is always right about every decision that he makes.  If you do not do it exactly the way he wants,  no matter how wrong he might be,  then you lose the contract and you don’t get paid.  And,  the government is quite often wrong about how best to do things!  That’s not to say the contractors are always right,  but they are wrong a lot less often than the government.

You also have to understand that NASA never did know,  and still does not know,  the art of building reliable solid propellant rockets.  Essentially,  no one at NASA ever did that kind of work.  They buy these things from contractors who (by definition) know much more of the science,  and especially the art,  than anyone at NASA knows.  The “science” is that knowledge which was written down.  The “art” is the knowledge that was not written down,  usually because no one wanted to pay for the writing.

I can tell you from experience as an insider within the business,  that “rocket science” isn’t really “science”!  It is only about 40% science,  about 50% art,  and about 10% blind dumb luck.  And that’s in production work!  In new product development work,  the art and luck percentages are even higher. 

Further,  this same sentiment applies to pretty much any type of engineering effort,  not just rocket work.  That explains a lot,  about a lot of things,  doesn’t it?

Poorly-Designed O-Ring Seal Joints

What I show in Figure 1 is how such joints should be designed and built.  This is the design that most solid rocket motors use,  very successfully,  whether large or small.  In most rocket motors,  you need only join the aft and forward closures to the case cylinder.  Only in some of the really large motors,  the case cylinder itself is divided into segments that must be joined,  usually to limit the size of the case-bonded propellant masses that must be cast and cured within them.

The sketch in the figure is what mechanical engineers call a “radial static seal”.  It is “radial” because the O-ring lies between an inner and an outer surface,  that must include a gap of tightly-controlled size between the two parts,  for assembly.  One part stabs into/inside the other,  in order to join them,  in this case by a row of pins.  It is “static”,  because the parts,  once joined,  do not move anymore.  There are strict but well-published guidelines and procedures for sizing the O-ring groove dimensions,  the gap for assembly,  and the size of the O-ring,  as well as its material composition and its hardness.  These guidelines and procedures are used precisely because they work so very well.   Examples:  Refs. 1 and 2.

Something also shown in the figure is peculiar to solid rocket motors,  especially those that are segmented-case designs.  There is a joint in the insulation (and thus also the propellant) that leads to the sealing surface gap,  that in turn leads directly to the O-ring in its groove.  You DO NOT obstruct this path with sealants,  putties,  greases,  or anything else!  But there does need to be a right-angle bend,  to stop radiant heat transfer from the flame in the motor from heating the O-ring directly.

The air in this path is what gets suddenly compressed upon motor pressurization,  and which in turn forces the O-ring to the far side of its groove,  where it gets squeezed against that surface to seal.  This is called “seating the O-ring”,  and until it is properly seated,  it CANNOT seal,  and so it briefly leaks!

Figure 1 – A Properly Designed O-Ring Seal Joint

It is the air in the path that gets compressed against the O-ring,  with hot booster gases and hot solids filling most of the path volume that the air formerly occupied.  But the air cools by convection to the steel much more effectively and faster than to the O-ring itself.  THAT is how the O-ring is not damaged by the hot air,  or the hot gases!  The hot solids are stopped by the right-angle bend.  This is a rapid transient on a time scale equal to,  or shorter than,  the motor pressurization event. 

What you DO NOT want is contact of the hot gases (and especially the hot solids) upon the O-ring!  The “hot sandblast” effect of that outcome would cut through the O-ring almost instantaneously.

Note that these two design requirements of (1) one O-ring and (2) an unobstructed pressurization path,  will interact very strongly with how one verifies proper assembly of the motor!  You must do a pressure leak check of the motor to verify sealing,  but you must do it by pressurizing the entire motor!  However,  you NEED NOT pressurize the motor to its full operating pressure to do this verification! 

You only need an atmosphere or so of pressure difference to seat any O-ring and then verify its sealing.  If it holds at that low pressure,  and you followed the design guidelines correctly,  it will hold at full motor operating pressure!  THAT is what you verify when you do motor case hydroburst testing,  long before you ever cast propellant to make a live motor!  That’s the way the real solid rocket motor manufacturers prefer to do it.  And it works to very high reliability levels,  as indicated in the figure.

However,  that is NOT what NASA insisted upon doing!  In the mistaken belief that a second back-up O-ring increases sealing reliability,  they insisted upon the two O-ring design indicated in Figure 2.  Thiokol complied,  lest they lose the contract.  In the mistaken belief that they had to pressure leak check at full motor operating pressure,  NASA did not want to risk fully pressurizing a live loaded motor (and rightly so).  And so NASA insisted on a way to apply air pressure at full motor operating pressure,  between each pair of O-rings at every joint,  instead of any motor pressurization.   This is shown in the figure.

What this does is drive the downstream (backup) O-ring to the correct side of the groove,  thus seating it for motor operation.  But,  it also drives the upstream (primary) O-ring to the wrong side of its groove,  from which motor pressurization upon ignition must unseat it,  drive it across its groove,  and re-seat it on the correct side!  Until and unless it re-seats on that correct side,  the upstream (primary) seal ALWAYS leaks!  Period!  There is NO WAY AROUND that outcome!  And THAT lets hot gases and solids reach the primary O-ring,  simply because the re-seating process takes a longer time than pressurization!

Figure 2 – The Improperly-Designed 2 O-ring Joint That Flew,  Up Through Challenger

NASA made a third mistake:  in the mistaken belief that it would prevent hot gases and hot solids from reaching the O-ring,  they insisted on obstructing the pressurization path by filling the insulation joint with “heat protective” putty (zinc chromate putty actually).  This is also shown in the figure.

This last mistake makes a bad risk even far worse,  because high pressure gases always (ALWAYS!!!) “wormhole-through” a not-solid material (like putty or grease) at a single point!  THIS effect is also shown in the figure.  That re-distributes the “push” of the gas from a broad front all around the O-ring,  to a single point upon the O-ring,  as indicated in the figure.  The delay unseating the ring,  pushing it to the other side of the groove,  and reseating it,  almost guarantees that the compressed air leaks past it,  so that booster hot gases and solids can reach the O-ring.  And those will cut a hole right through it.

“Half-moon slices” right through the primary upstream O-ring were seen,  upon SRB motor disassembly,  in a rather significant percentage of the SRB’s recovered and refurbished.  That verifies what I just said about the upstream O-ring being cut!  There is no surprise there,  once you understand the process!

The difference between this point load problem,  and what NASA analyzed in its structural calculations for the O-ring seal is quite stark!  The structural analysts were assuming pressurization on a broad front.  They did not model the point load effect of the hot gases and solids wormholing-through the putty obstructing the pressurization path.  Quite simply,  what was built was NOT what was analyzed!

Unnecessary Risk to Fly Too Cold

If the motor is sufficiently cold-soaked,  the primary upstream O-ring loses its flexibility and resilience (as do all of them).  Pushing the entire embrittled O-ring across its groove all at once is risky enough,  but if you concentrate the “push” at one single location by the wormhole effect,  you essentially guarantee snapping the O-ring apart at that point!  This cold brittleness effect was amply demonstrated by Dr. Feynman at the Rogers Commission hearings (assisted by Gen. Kutyna),  when he stirred his sample of the O-ring material in his glass of ice water,  and then demonstrated its non-resilience.

Any failure of the primary upstream O-ring,  whether by hot sandblast cutting,  or by cold brittle fracture from the point jet force load,  then puts a single-point hot sandblast jet impacting onto the downstream O-ring,  simply because it is nearby!  Thus,  a sort of “cascade failure” is a very high risk indeed!

The post-Challenger “fix” was a third O-ring in every joint.  This just set up the cascade failure as a longer chain,  as indicated in Figure 3.  The only reason the Challenger disaster did not repeat is that they never flew that cold again.  But the 1/51 failure rate demonstrated by loss of Challenger speaks for itself!

Figure 3 – The Cascade Failure Risk Was Compounded By the Redesigned Joint

Fatal Consequences We All Saw

The photography obtained during the launch and loss of Challenger confirms everything claimed here.  The seal failed upon motor ignition and pressurization,  as shown quite clearly in Figure 4.  The dark grey plume is carbon soot-bearing hot gases spewing through the two failed O-rings at the aft segment joint. 

Figure 4 – Seal Leak Upon Ignition Seen In Photography

This leak miraculously “cured” itself by plugging-up with aluminum oxide-carbon slag from the metallized propellant.  This slag-plugging just happened to hold pressure like that,  until the Challenger encountered a wind shear while at “max-Q”,  where it was also most highly stressed by aerodynamic forces.  The slag plug failed,  letting the hot motor gases and solids rush through the hole again.  This is shown quite clearly as the anomalous bright-but-small extra plume in Figure 5 below.

This jet of leaking hot gases and solids finally got so big that it cut through one of the aft struts holding the SRB to the center tank.  There is always hydrogen leaking from the center tank’s hydrogen tank,  and in this case the leaked plume probably burned a hole in that hydrogen tank.  With the strut cut,  the bottom of the SRB moved outboard.  That pushed the nose of the SRB inboard,  such that the nose of the SRB poked a hole in the side of the center tank’s oxygen tank. 

Suddenly dumping oxygen into a base-burning hydrogen-air fire caused an explosion in the wake behind the center tank that both overheated and structurally overloaded it.  The tank collapsed,  letting both SRB’s and the orbiter fly free.   The released propellants burned explosively as this happened.  All this happened in an instant,  so it looks like just the one sudden explosion.

Figure 5 – Leakage Resumed After Being Shaken By Wind Shear at Max Q

The released SRB’s continued to “fly” out-of-control under their own thrusts,  as we all saw.  This is shown in Figure 6.  The orbiter’s engines were pointed through a center of gravity that suddenly no longer existed,  so they forced the orbiter to pitch-up violently,  before starving for lack of propellant from the suddenly-missing center tank.  The pitched-up orbiter went broadside to the supersonic wind,  which tore it to pieces.  This is how those pieces,  that we all saw fall into the sea,  came to be.  

Figure 6 – The SRB Did Not Explode,  But It Punched a Hole In the Center Tank

Final Remarks

The two O-ring joint was a NASA-mandated design mistake,  compounded by mandating putty obstructing the O-ring pressurization paths.  The “customer is always right” in government contracting,  except that he was lethally and fatally wrong about this one!  See also Ref. 3.

The decision to fly cold-soaked colder than the SRB’s had ever been tested,  was also a NASA management decision.  Both NASA and Thiokol engineers objected,  but were over-ruled.  Thiokol upper management also over-ruled their own engineers,  and told NASA to go ahead and launch.  Thus emboldened by Thiokol management,  NASA launched the thing,  thus killing its crew.

The stand-down to “correct” this problem was nearly 2 years long and horribly expensive.  Which just goes to prove what I like to say to anyone who will listen:  “there is nothing as expensive as a dead crew,  especially one dead from a bad management decision”. 

The only problem with that return-to-flight effort is that they did not correct the real problems upon return-to-flight,  they actually made them worse with a 3-O-ring joint,  and by keeping the putty obstructions.  The ONLY thing they did “right” was never to fly that cold again!  Which is very likely the ONLY reason that the Challenger disaster did not repeat itself before the Shuttle got retired,  since there were more than 51 more flights after the Challenger disaster!

By the way,  the crew did not die in the tank explosion and subsequent ripping-apart of the orbiter by air loads.  The telemetry showed no high-gee accelerations at all!  The crew was still alive in the orbiter cabin until it finally hit the sea,  which is about a 200-gee stop,  since it hit dead broadside.  See Figure 7.

Figure 7 – The Crew Was Still Alive In This Cabin Section (Arrow) That Is Falling Back

I say what I said about the crew because the flight deck back-seaters leaned forward and flipped on the breathing-air packs for the front-seater pilots.  They would not have done that unless they knew the cabin had depressurized,  and that would have been significantly AFTER the explosion and ripping-apart of the orbiter.  They were tumbling clear of the explosion cloud by that time,  as illustrated in the figure. 

Those two flight deck pilots had breathed-up all the oxygen in their breathing packs by the time they hit the sea,  something confirmed by the empty breathing packs that were recovered.  Which means they were alive when they hit the sea!  By extension,  so were the back-seaters,  plus the three down on the mid-deck.

They did not have pressure suits,  parachutes,  breathing bottles,  and a hatch they could blow open (basic bail-out gear).  More importantly,  there was no way to take the spin off the tumbling cabin.  Spinning like that,  there was no way to reach and exit the hatch,  even if they had the other basic bailout gear!  But a small drogue parachute from the nose of the cabin section would have taken off the spin!  That plus the basic bail-out gear just listed could have saved that crew!  It took almost 5 minutes to hit the sea.  They had the time to bail out.

I submitted that means for bail-out to NASA,  but I was ignored.  Coming from an outsider,  my idea was “not invented here”,  as far as NASA was concerned.  Yet,  something rather like it might even have worked for Columbia some years later:  the 3 mid-deck occupants were still alive inside a tumbling cabin section as it approached impact near Tyler,  Texas,  well after the breakup during re-entry.   Time was short for a bail-out,  but without the de-spin drogue,  they could not reach the hatch at all. 

References

#1. Parker O-ring Handbook ORD 5700,  copyright 2021,  original release 1957,  Parker O-Ring and Engineered Seals Division,  Lexington,  KY, available from parkerorings.com 

#2. Seal Design Guide,  Apple Rubber Products,  Lancaster NY,  available from AppleRubber.com

#3. Wikipedia article “Rogers Commission Report”,  in this case accessed 11-26-2021

Final Notes

There are different design rules for static radial and static face seals,  and different rules yet for dynamic radial seals (as on a piston moving inside a cylinder,  like a syringe or a hydraulic cylinder).  The Shuttle SRB joints fall into the static radial classification. 

The appropriate set of rules specifies O-ring sizes and hardness,  groove dimensions,  and when to use back-up rings.  You just follow the design rules,  and make sure that only compressed air reaches the O-ring (and on a broad front),  upon solid propellant motor ignition.  

You accomplish that broad-front pressurization with the 90-degree bend geometry to stop the hot solids and radiant heat transfer,  and by NEVER obstructing the O-ring pressurization path with anything!  Even too close a fit between the hard parts,  can cause problems with the transient pressurizing flow.

You verify your seal design,  your case structural design,  and your leak check procedure,  during case hydroburst testing,  long before you ever cast a live motor!  You NEVER delete the hydroburst testing step in your development effort.  Never!  Not for any reason at all! 

Then you test live motors at every environmental extreme condition in which you think you might possibly operate.  If any redesigns (of anything) are needed,  you go back and verify them in all the tests,  from hydroburst all the way forward.  No motor goes to production,  until its exact design configuration has been verified in every test at every test condition!

Once your design has passed all those tests,  you stick with your verified leak check procedure as if it were a religious mandate!  You add rigorous quality control (of the “total quality management” type),  for production.  That includes X-raying every single item,  to verify that there are no casting voids in the propellant,  no unbonds between propellant and case liner,  and no other propellant grain cracks or other problems.  And then you NEVER operate a motor outside the conditions for which it was tested! 

THAT is the way to achieve no-more-than-1-in-a-million failure rates,  with solid propellant rocket motors!

The “bean counters” and “management professionals” will absolutely hate that prescription as “too expensive”,  but killing a crew with a bad design just costs a whole lot more,  than the cost of following that prescription.  We’ve already seen that with Apollo-1,  Challenger,  and Columbia.

Simple as that. 

And just as hard to sell to the “bean counters” and “management professionals”,  as you might fear.  

The “Warm Brick” Ramjet Device

One of the most unusual ramjet projects I ever worked on was a non-propulsive device.   This was a very miniature ram-fed airbreathing combustor,  that was to be the hot gas generator for an infrared (IR) decoy.  This decoy was to be towed behind an aircraft in lieu of a whole series of dispensed flares.  It was intended to work by having enough IR output to cause the aircraft to drop out of the missile field-of-view first.  See Figure 1 for that concept.  I was working for my friend Byron Hinderer doing this.

Figure 1 – Towed IR Decoy Concept,  called “Warm Brick” at Tracor in 1984

I did this at what was then Tracor Aerospace,  in Austin,  Texas,  during 1984.  We called this decoy “Warm Brick”,  and my job was to determine if this concept was even feasible (it was).  Our idea was to heat a porous refractory material until it glowed brightly in the IR.  We preferred fuel-air combustion to minimize decoy mass,  and ram combustion is the simplest of the airbreathers.  Plus,  I had lots of experience with ramjet combustion at what was then Rocketdyne/Hercules in McGregor,  Texas.

To the very best of my knowledge,  no patent was ever taken out on this concept,  and Tracor never did anything at all with it.  Even if there had been a patent,  and it had been renewed,  any such patent would have run out by now.   So,  what I reveal here should offend no one,  and infringe no patents.

As implemented for the feasibility tests,  this concept took the form of a “gasoline lantern mantle” made out of commercial ceramic fire curtain cloth,  as the IR emitter.  This was to be mounted behind a wake-producing spoiler,  mounted at the aft end of the burner and inlet assembly.  The decoy might carry its own fuel tank,  or it receive fuel down its tow line,  if a heavier tow line could be tolerated. 

To test the scientific and engineering feasibility,  I designed a very generalized inlet and burner hardware set that was flexible enough to allow evaluation with a variety of gaseous and liquid fuels.  See Figures 2 and 3.  The intended flight conditions were relatively low altitude from mild subsonic to barely-supersonic speeds,  typical of an attack aircraft threatened by surface-to-air missiles.  

Figure 2 – Assembly Sketch for the Initial Version “Warm Brick” Ram Combustor Test Device

The assembly sketch clearly depicts the long fuel injection-and-mixing duct allowed between the inlet diffuser and the sudden dump into the combustor.  There was an inlet piece and a fuel injector piece,  both made of aluminum for ease of rework,  and an inlet tube and a combustor shell,  both made of steel.  The combustor shell was sized for fabrication from 2-inch schedule-40 pipe,  but ended up being made of 300-series stainless to those same dimensions.  We tried automotive-style spark ignition.

One can easily see how the molded low-density ceramic liner insert was to be trapped in place by the nozzle block.  The arrangement shown in the assembly sketch of Figure 2 (directly-pinned nozzle block) was quickly replaced by a pinned steel nozzle shell ring,  as shown in the hardware photo (Figure 3).  This revision happened about the same time that the first (unreinforced) liner was replaced with the second liner (reinforced ceramic composite).  

Figure 3 – Photo of the “Warm Brick” Ram Combustor Test Hardware as Revised

The design concept called for a small combustor fed by a simple pitot inlet,  with a convergent-only nozzle that would likely function unchoked at most conditions.  I chose a center-duct coaxial air entry with sudden-dump flame stabilization,  similar to the successfully-flown ASALM-PTV liquid-fueled ramjet test vehicle.   Geometric ratios were initially set equal to those used in ASALM.

Based on Reference 1,  I chose a minimum ¼-inch (6 mm) step height around the dump.  The combustor length was sized “empirically” (rules of thumb based on ASALM-PTV geometry) so that the annular separation bubbles would close,  and the axial core would be “burned out”,  before any of these flows entered the nozzle.  That was basically an assumed 11-degree spreading angle,  on both sides of the mixing layer between the entering mixture and the recirculated flame.   That’s too crude,  in hindsight.

We wanted sufficient porosity in the emitter so that the burner operation would be unaffected by the presence or absence of the emitter.  The fire curtain cloth gave us that,  in the sizes tested,  because the surface area of the ellipsoidal shape was so large relative to the final burner throat area.  Its effective porosity-driven “free” open area was very much larger than any of the burner throat areas we tested.   

There were two crucial unanswered questions:  (1) emitter/hot gas coupling (could we really get the emitter hot enough to radiate effectively?),  and (2) obtaining stable combustion at all in a burner that small,  with any fuel whatsoever!  There was an extensive paper trade study done,  to determine the desired fuels.  In test,  these fuels,  and some other fuels that were easier to use,  were investigated. 

This combustor was nominally 1.5 inch (38 mm) inside diameter,  as insulated,  and 3 inches (76 mm) long inside.  The smallest size ramjet combustor in my experience up to that point had been some heavyweight solid-propellant ducted rocket ramjet work (in a completely-different geometry) at 4.6 inch (117 mm) inside diameter,  and length/diameter 6-to-8.  The largest was ASALM-PTV at a 20 inch (51 cm) combustor case diameter.  “Warm Brick” was smaller than anything of which I had any knowledge! 

I didn’t want to periodically replace an ablative liner in the test burner,  and I didn’t want to attempt an air-cooled liner shell for full-rich combustion in something that small.  So I opted for an unknown,  inspired by the Space Shuttle’s heat shield tiles.  Could I put a low-density ceramic insulator in this combustor,  and not melt it?  The answer turned out to be “yes”,  but it took some adaptive effort.

The project operated in three logical parts:  (1) obtain stable combustion with a variety of fuels in the burner alone,  (2) add the emitter and determine how best to shape,  fabricate,  and attach it,  and (3) document infrared radiometric output.  The real prerequisite for part (1) was the combustor insulator,  since we started with gaseous fuels,  thereby avoiding the fuel vaporization issue. 

I selected free-jet test mode as the best way to accomplish all three parts of this project with the same hardware and test setup (see Figure 4).   All that I had personally done while at Rocketdyne/Hercules was direct-connect testing,  but I knew about free-jet testing,  both from my research,   and some experimental association with Marquardt,  while I was with Rocketdyne/Hercules. 

We used a commercially-rented air compressor trailer as our air source,  to be run real-time.  In 1984,  this 750 SCFM unit was the largest of its kind in Texas.  It fed a PVC pipe stilling chamber,  terminating in a simple convergent-only nozzle block made (conveniently) of wood.  

Figure 4 – Test Setup:  Stilling Chamber Exhausting To Left,  Fed From Right

The test article was bolted to a heavy pipe stand-and-sting,  with its inlet immersed in the free jet of air.  That free jet typically measured 190 F (88 C) stagnation temperature,  at full-power compression conditions. 

The first part of the investigation began with bottled hydrogen gas fuel (series 1).  This and all the other trials are summarized in Table 1 below.  Series 1 wasn’t very successful for two reasons:  (1) the nozzle was too wide open for a stable flame,  and (2) free jet air speeds higher than about 0.25 Mach blew the spark column out from the electrodes of the spark plug,  even though it was located flush within the annular recirculation zone.  

The device didn’t ignite at all until I obstructed the nozzle with a scrap of wood,  and it still went out after ignition,  if I removed the obstruction.  So,  I built a smaller-throat nozzle block.  We still had to ignite at low airspeed and gradually work up to higher speeds,  limited at that time to about half a Mach number by the stilling chamber nozzle.  I also tried liquid ethanol unsuccessfully at this time (series 2). 

Somewhere in all of this,  I first drove the combustor into what proved to be a very violent rich blowout instability,  and completely shattered my first (unreinforced) liner!  The combustor visibly shook on its sting,  and it spit the pieces of its liner out the nozzle,  igniting a local grass fire!  Later,  we estimated a pressure amplitude near 0.8 atm,  at audio frequencies (a few hundred Hertz),  for this instability. 

A photo of the liner molding tools that I used is given in Figure 5,  which includes the basic combustor shell as the outer forming tool for the combustor liner.  Both it and the nozzle block were laid up as (commercial) low-density molding compound troweled onto the wooden plug,  and inserted into the corresponding shell for cure.  I used Cotronics Corp. 360M low-density molding compound for this.  

Figure 5 – Tools Used for Molding Ceramic Combustor Insulation Liner Inserts

These parts were cured at 215 F (102 C) in an oven to drive off the water.  The circuitous paths for exiting steam led to a low density ceramic matrix.  The resulting parts were coated with a paint-like ceramic cement (Cotronics Corp. 901),  and cured again,  in the same oven.  The unreinforced liner showed evidence of hot gas flow behind the insulation,  and into the cracks,  shown in Figures 6 and 7.  

Figure 6 – Recovered Pieces of Shattered Unreinforced Liner,  Bonded Together

Figure 7 – View of Fracture Surface,  Showing Hot Gas Flow Damage with Sooting

I built a second ceramic composite liner reinforced by layers of the fire curtain cloth (woven from 3M Nextel 312),  which survived all instabilities and any other test abuses thereafter.  It survived many hours of accumulated burn time in near-pristine condition,  as seen in Figures 8 and 9.  The shrinkage cracks did not preclude functionality.  There was some melting evident in the throat of the nozzle.  

Figure 8 – View Into Near-Pristine Reinforced Liner,  After Hours of Burn Time

Figure 9 – View Into Reinforced Nozzle Block,  After Hours of Burn Time

Once we had the burner working at all,  we tried some test sample pyrometers in its exhaust plume,  with both propane and acetone as fuel (series 3,  and acetone proved worthless as a fuel).  These pyrometers would be old nails,  or else planar samples of potential emitter materials.  We even tried gasoline as fuel (series 4),  but results were poor,  and it became very obvious that poor vaporization was the cause!  I tried propane again (series 5) as the most successful fuel,  and got enough radiometer output to be encouraging,  from a sample of the fire curtain cloth immersed in the jet exhaust. 

So,  I created a fuel-line hot-soak bucket to correct the poor fuel vaporization problem for test purposes.  This took the form of an electrically-heated bucket of old motor oil,  in which a coil of the fuel supply line was immersed.  That rig is shown in Figure 10.  It may resemble a moonshine still,  but it is not!  

Figure 10 – Fuel Vaporization Preheat Bucket Rig

At this point,  I had a crudely-successful burner,  but an unproven fuel supply method.  I checked out the combined burner and fuel vaporization bucket,  first on propane (series 6),  then on aviation gasoline (series 7),  and finally on a “home-made version of JP-4”  that was actually half Jet-A and half aviation gasoline (series 8).  Plus,  I added instrumentation to the burner unit (enough manometer pressures and thermocouples to attempt an actual “engine” cycle analysis). 

Results,  including the exhaust pyrometer samples,  were favorable enough to warrant continuing the project further.  It still required a lower-airspeed ignition.  I stood in the jet blast for all these tests,  looking directly into the flame zone,  and sniffing for unburned fuel,  to set mixture.  That “settled” the fuel injection and ignition issues well enough to test emitter coupling issues for the very first time. 

The first actual emitter was made of Nextel 312 fire curtain cloth,  coated with the Cotronics 901 adhesive as a “paint”.  It was sewn together,  with alumino-silicate thread,  from bias-cut gores much like a balloon,  to form an elongated semi-ellipse approximation.  The seams were left on the outside of this first emitter,  as shown in Figure 11.  It was the first of several series 9 tests with pre-heated propane,  at air speeds up to about Mach 0.47.  Those test conditions are depicted in Figure 12.

Figure 11 – Test Setup for First-Article Emitter

For all subsequent tests,  the seams in the sewn emitters were placed to the inside,  as is depicted in Figure 13.  That photo shows post-burn appearance of two series 12 emitters tested with ethanol fuel,  but all the internal-seam emitters appeared similar,  regardless of series and fuel. 

These articles were brittle and fragile post-test,  as expected for alumino-silicate materials soaked to temperatures exceeding the solid phase-change temperature of about 2350 F (1290 C).  That fragility alone confirmed a high surface temperature for radiation purposes!  This was also verified by radiometric measurement,  which also indicated very “non-gray” behavior,  in that the effective color temperature (radiation peak wavelength) was substantially cooler than the actual temperature. 

The spoiler just ahead of the emitter clamp mounting provided protection from direct wind blast forces.  Plus,  it also provided effective hot gas recirculation effects external to the emitter surface.  Both acted to raise emitter material soak temperature,  and therefore IR output,  quite successfully.    

Figure 12 – Test Conditions Explored with Series 9 Propane

Two tests were made as series 10 in this same configuration with the “home-made JP-4” fuel.  Results were similar to the series 9 propane runs,  except for a small liquid-wet “cold spot” at the very end of the emitter bulb.  This was due to still-unvaporized kerosene hitting the emitter on-axis.  

Figure 13 – Post-Test Emitter Appearance from Series 12 Ethanol Tests

Sometime during this checkout process before the series 9 propane runs,  I successfully modified the inlet to a larger lip radius, in order to decrease its “buzz” instability tendencies at higher backpressures.  That also greatly improved ignition characteristics,  and it further pushed the rich blow-out instability limits to richer mixtures!  The test set-up for cold-flow inlet calibration is shown in Figure 14. 

Both the original and modified (larger lip radius) inlets were cold-flow tested with this rig.  Data were cross-plotted in a variety of ways.  The data plot format for “typical” supersonic ramjets was rather undiscriminating at these subsonic speeds:  stream tube area ratio versus Mach and stagnation pressure recovery ratio versus Mach.  Plots in the more primitive-variable format were actually more useful for this mostly-subsonic system.  These included the diffused Mach to freestream Mach ratio,  and the static “pressure gain” ratio. 

These results guided the 1984-vintage data reductions of the series 9 propane runs with emitters.  From those,  installed hot-burn test inlet performance data matched the cold-flow tests.  The streamtube area recovery ratio shows a very strong influence of the so-called “highlight” area versus the true minimum area,  when used as the reference area for the calculation. 

After the fact,  this was entirely expected,  based on Reference 2,  which (of course) recommends the highlight definition.  At the time I did these tests,  I had used something pretty close to the minimum area for the reference.   It shows explicitly in the data,  as a recovery ratio substantially greater than unity,  which is completely out-of-line with the usual expectations for ramjet inlets. 

See Figures 15 and 16.  

Figure 14 – Cold-Flow Inlet Calibration Test Rig

After the series 9 and 10 tests,  the air nozzle in the stilling chamber was replaced with a second wooden unit of slightly smaller throat diameter,  as depicted in Figure 17.  This enabled free jets of nearly Mach 1 speed at the maximum compressor output.  Two more test series were conducted with this change,  specifically to obtain data at those higher simulated air speeds.  These were series 11,  using both propane and hydrogen fuels,  and series 12,  which used the finally-selected ethanol fuel. 

The series 12 tests employed both radiometer measurements,  and imaging with a thermal imager camera.  The fuel vaporizer rig was less successful with a high latent heat pure-substance fuel (ethanol),  than it had been with distillate fuels,  or with the easily-vaporized propane.  With ethanol,  it was essentially long-period unstable,  with an oscillating fuel flow output.  The cycling time was a few seconds. 

Nevertheless,  using ethanol fuel produced an output spectral power distribution closer to what is needed from the non-gray decoy.  The radiometer data clearly showed this.  We attributed this difference (with a high degree of confidence) to the lack of yellow carbon glare in the ethanol flame.  This yellow carbon glare was quite noticeable in the propane tests,  and even more so when using gasoline or jet fuel.  The series 12 ethanol runs looked to the eye “positively white” in comparison.

The ethanol fuel injector was stopwatch-and-bucket calibrated for those series 12 tests. Those calibration data are shown in Figure 18.

Figure 15 – Calibrated Inlet Performance Derived from Series 9 Data,  Part 1

Figure 16 -- Calibrated Inlet Performance Derived from Series 9 Data,  Part 2

Figure 17 – Air Nozzle Re-Work for Higher-Airspeed Test Capability

Figure 18 – Flow Calibration Data for the Series 12 Ethanol Fuel Runs

After these tests,  the fuel vaporization problem was conceptually addressed as a hot-gas tap from the forward end of the combustor to the lower-pressure zone at the minimum area of the inlet.  Fuel would be injected into this very hot recirculated gas stream to effect rapid vaporization.  While the design analysis looked good,  that concept never received any testing due to budgetary constraints that essentially stopped all experimental work on the project after late 1984.  Some prototype flyable hardware was designed,  and a few of those parts manufactured,  before all work on the project was completely stopped.  It never resumed.  So NOTHING is confirmed about any of this!

The ceramic liner material was never characterized,  it “just worked”.  Density,  strength,  and thermal conductivity were never measured in any way!  However,  it handled as if it were about as dense as commercial Styrofoam products.  The strength was considerable,  considering the rich blow-out instability abuse it endured.  Immersed in a 190 F (88 C) air stream,  the combustor shell would “barely boil spit” after an hour-long burn test at full rich mixture (theoretically around 3800 F or 2100 C),  with but 0.2 inch (5 mm) thickness of the insulation!  That indicated very low thermal conductivity indeed! 

Table 1 – Summary of “Warm Brick” Burner Tests

In recent years,  I developed further those basic cycle analysis techniques applicable to a low-speed ramjet system,  or a subsonic nonpropulsive item like “Warm Brick”.  In particular,  I programmed them into an “Excel” spreadsheet,  and reanalyzed the “typical” series 9 propane run at 0.47 Mach air speed and full-rich mixture.  The spreadsheet setup is shown in Figure 19,  and the spreadsheet results in Figure 20.  Since then,  I have created a real low-speed ramjet cycle analysis code.  It works just fine. 

These recent compressible-flow cycle analysis results defined the bulk flow conditions inside the combustor well enough to attempt a heat transfer model with a reasonable expectation of success.  That model was cylindrical convective-conductive,  and based on standard compressible flow models inside and outside the combustor shell.  Radiative loss was near zero,  as there was no effective path by which thermal radiation could leave the interior.  The shell radiation cooling potential was very low. 

While the steel shell has a well-known thermal conductivity,  the ceramic composite liner did not,  so I ran this model parametrically versus conductivity values from “very low” to “very high”.  The “best” value of thermal conductivity was that which matched both my recollections of perceived shell temperature,  and my observation that the liner surface was often close to melting (3250 F, 1790 C). 

Those thermal conductivity results are given in Figures 21 and 22.  The highlighted value of 0.02 BTU/hr-ft-F equates to 0.035 W/m-C.  Density and strength still lack actual characterization!  I have often wondered whether this material might serve as a re-entry heat shield material,  the way that the somewhat-similar low-density ceramic Shuttle tile did.  But that is another topic for another venue. 

References:

#1. Curran,  Edward T.,  “An Investigation Of Flame Stability In A Coaxial Dump Combustor” (dissertation,  AFAPL/RJ WPAFB,  Dayton,  OH),  AFIT/AE/DS 79-1,  Jan. 1979. 

#2. Seddon, J.,  and Goldsmith,  E. L.,  “Intake Aerodynamics”,  AIAA Education Series,  1985,  ISBN 0-930403-03-7.

Figure 19 – Spreadsheet Setup for “Warm Brick” Cycle Analysis at Series 9 Propane Conditions

Figure 20 – Spreadsheet Cycle Analysis Results for “Warm Brick” at Series 9 Propane Conditions

Figure 21 – Heat Transfer Model Results for “Warm Brick” Liner Thermal Conductivity

Figure 22 – Heat Transfer Model Results Plotted vs Radius

Epilogue:  Some Practical Combustion Device Lessons Learned

Cycle analysis with one-dimensional flow models turned out to be less important than the actual scale-dependent physical chemistry of flame stability,  for this “Warm Brick”  device.  Residence time is proportional to dimension,  all else equal,  while chemical reaction rates are scale-independent.  This alone suggests there is a minimum size below which a thing “just won’t work” with a particular fuel. 

Mixing is another very strong determinant of flame stability.  Mixing is not proportional to scale,  nor is it scale-independent,  but it is something in-between.  Again,  this also suggests that there is a size below which a thing “just won’t work” with any particular fuel.  That is precisely one issue (of many) in flameholding.

Those considerations explain why the required nozzle contraction ratio (and internal flow velocities) were so low in the “Warm Brick” device for stable ignition and burning,  relative to everything I knew about,  before I attempted this project.  However,  these experiences with the Warm Brick subminiature combustor predate the in-depth understanding of flameholding and flame stability that I was later able to achieve,  after returning to Rocketdyne/Hercules.  That knowledge is summarized in the “exrocketman” article titled “Ramjet Flameholding” (on this site) and dated 3 March 2020.

The vaporization of fuels of different latent heats and boiling behavior revealed a surge instability in the hot-bucket fuel rig (referring again to the crude hardware in Fig. 10 above).  The basic layout was a source of fuel at pressure,  led through a copper line coiled in the hot bucket,  and from there to the metering orifices inside the test article.  See the cartoon in Fig. 23. 

The source of fuel-at-pressure was a standard 5-gallon propane bottle (usually around 200-250 psig),  or a welding gas bottle (initially 2200 psig),  or a pressure tank of liquid fuel pressurized with compressed dry nitrogen (usually pressurized in the 100-300 psig range).  All of these pressurization schemes are regulator-controlled.  That regulator was physically located about 5-to-10 feet downstream of the test article,  and within arm’s reach of the exhaust plume.  This allowed me to manually adjust the fuel flow during the test by varying the regulated pressures,  while standing immersed in the exhaust where I could smell for unburned fuel.  For the open-nozzle tests,  I could literally see the flame up the tailpipe.    

Fig. 23 – Conceptual Layout and Operation of Fuel Supply

When using hydrogen directly from the welding gas bottle,  there was no vaporization problem,  as this was simply compressed hydrogen gas.  We did not use a pre-heater bucket with this fuel,  but the rest of the component layout in Fig. 22 is correct. 

With propane in the 0.47 Mach air tests,  we found the line just downstream of the regulator,  and the sides of the propane bottle,  to be cold.  This is because the vaporizing pool of liquid propane in the bottle must draw about 150 BTU/lbm of latent heat from itself and from its surroundings,  mostly from itself (gets cold).  If it cannot draw sufficient heat to vaporize,  then it won’t vaporize,  pressure drop notwithstanding!  The energy to change phase (latent heat) simply must come from somewhere!

There was a cold-line risk of re-condensation on the way to the test article,  which we “cured” with the hot oil bucket preheater.  We kept the line length from bottle to preheater as short as practical.  We also found bottle “freeze-up” occurred at the higher flow rates with the Mach 0.9 airstream tests.  We “cured” that by the camper’s expedient of putting the propane bottle in a tub of hot water. 

With gasoline and jet fuel,  the driving pressures helped us pre-heat the liquid fuel without getting any boiling in the fuel line.  Without preheat,  there was insufficient air stream heat in the test article to get the fuel to vaporize and burn.  With about 300 F preheat,  we got all but the “tag-end” of the distillation curve to vaporize upon being injected,  due to combined atomization and pressure-drop boiling. 

With the gasoline and 300 F preheat,  our nominal 100-300 psig driving pressure was apparently barely enough to prevent any significant boiling in the line,  so we did not encounter any noticeable problems with vapor lock-induced fuel flow rate surges.  With the jet fuel and its lower volatility,  we had no real risk of vapor lock surging,  but we did see a little more “tag end” unvaporized fuel,  indicating a higher preheat temperature was really needed.  Both of these are about 150 BTU/lbm latent heat materials.

We did have a real fuel surge problem running neat ethanol as fuel.  This material has a far higher latent heat at about 378 BTU/lbm,  and it has a single normal boiling point,  instead of distillation behavior.  At our delivery conditions,  the pressure was insufficient to prevent boiling in the line,  leading directly to vapor lock-induced flow rate surging!  Fuel delivery rates oscillated through about a factor of two,  on a long period of several seconds.  It would vapor lock,  unlock,  and relock to cause this surging.

We could not reduce preheat temperatures and still expect to get any flash vaporization upon injection,  in hindsight due to that higher latent heat.   We could not increase the feed pressures to preclude the boiling without re-working the test article for much smaller injection orifices.  That latter is the real design solution to this problem,  but we did not use it for these tests!  We were able to get our infrared radiometer data from the high points of the oscillating-intensity burn. 

While high pressure preheat to get flash vaporization from an atomizing injector is an approach that really works,  the equipment to do it is usually large and heavy,  too much so for a miniature decoy.  The alternative would be to mix the fuel with hot combustion gas to get vaporization,  downstream of the metering point.  The design difficulty is then to get good mixing of the fuel-rich gas stream with the inlet airstream,  without suffering large pressure losses.  That seemed the better approach for the flight decoy design.  We were never able to test this,  though!  It is still just a concept!

For an aero-engine application,  high-pressure fuel pre-heat with atomizing flash vaporization is likely the better design approach.  The sizing of required preheat depends upon raising the liquid to a temperature such that the enthalpy drop across the injection orifice exceeds the latent heat of vaporization.  The size of the orifice and the feed line pressure determine flow rate.  But,  the feed line pressures must always exceed fuel vapor pressure at that high preheat temperature!  If this is not done,  then vapor lock-induced surging will occur,  and at very significant magnitudes.  Fuel control then becomes impossible. 

As indicated,  we never got to test the concept of vaporization by injection into a hot combustion gas stream,  followed by injecting that hot mixed stream into inlet air.  There is a lot of promise in that notion,  but it is fraught with practical difficulties,  as well.

Final Comments:  IR Emission Characteristics and Towed Decoy Physics

The IR emission characteristics topic has been mostly ignored here,  except to say these ceramics were decidedly “non-gray” in their spectrally-dependent emissivity properties.  They were non-gray enough to reduce expected radiation in the 1-2 micron band very markedly,  to near what they emitted in the 3-5 micron band,  despite operating at a temperature somewhere near 3000 F (1650 C).  The effective “color temperature” (really the wavelength at peak spectral distribution of radiation) was much closer to typical tailpipe temperatures at full power (but less than those with full afterburning).

Suffice it to say that a great deal of infrared power was radiated by a very small object,  whose color temperature and radiated-power in-band looked like a very large jet engine tailpipe at full power.  This little emitter would blister my face with radiated heat from some 20 feet away.  The large radiated power would be the temperature-to-fourth-power effect,  while the color temperature would be the non-gray emissivity effect.  Both are critical effects.

Exploring this IR emission topic in more detail would be the subject of some future article,  or perhaps even a book relating these experiences.  This is an application of otherwise well-established physics.

Another unaddressed topic is aeromechanical in nature:   how to tow hard-body decoys stably on towlines,  at speeds from very subsonic to low supersonic.  The answers are not what one would expect,  based on the towed gunnery targets that have been flown for some decades now.  Straight tow is the easiest to achieve at all speeds,  meaning the tow line extends mostly straight back from the aircraft,  although you DO NOT tow the body by its nose! Low or high tow are far,  far more difficult to achieve,  especially as speeds become high subsonic and the aero forces exceed the weight force.  Stable side tow is nearly impossible,  even at low subsonic.  This applies to radar decoys as well as IR decoys.

Exploring how to tow hard bodies behind aircraft might be the topic of a future article or articles,  or even part of a book.  The basic rules were invented by my friend Byron Hinderer.  I researched the details,  and documented what did not work,  as well as what did,  in the wind tunnel while at Tracor.

The final unaddressed topic deals with what is called “engagement analysis”,  where the geometry of the aircraft,  the tow,  the approach geometry of the attacking aircraft or missile,  and the characteristics of the decoy (IR or radar) and the seeker,  all interplay.  The desired result is an estimate of the kill probability for the attack.  The decoy designer wishes to reduce that kill probability to near zero.

Exploring engagement analysis with IR decoys and IR threats might be some future article.  Or it might also be part of a book on these experiences.   This topic I learned and practiced while at Tracor.