“Entry By Hand”

Why Know This Stuff?

(1        (1)    It provides a more efficient way to expend limited resources (see Figure 1).

(2) It is integral to brainstorming,  raising probability of success with multiple ideas.

(3) No organization can afford to do real design work on all candidates!

(4) However,  it requires “real” pencil-&-paper engineering training.

    (5) Those so-capable can spot bad results coming from computer codes! 

Figure 1 – Knowing “Pencil-and-Paper Engineering” Is the Efficient Way to Use Resources

The example here is making entry estimates that include both dynamics and heating.  The basic by-hand entry model is old,  simple stuff used for warhead entry work back in the early 1950’s.  It is usually attributed to H. Julian Allen,  although both he and A. J. Eggers published it in a NACA report,  once this method was declassified.  See Figure 2.

This kind of analysis is now best done in spreadsheet, for fast changes,  that automate any iterative explorations.  This analysis only handles straight-in entries:  no skips, no multi-pass trajectories.  It is fundamentally 2-D Cartesian,  so one must “wrap” the range-related results around the central body.

Figure 2 – How the Old Entry Model of Allen and Eggers Actually Works

 The inputs divide into 3 groups:  (1) the atmosphere scale-height density model and entry interface altitude,  (2) the entry speed and direction information,  and (3) the vehicle model.  That last includes both ballistic coefficient as well as an effective “nose” radius for heating estimates.  Ballistic coefficient requires mass and dimensional information,  plus an estimate of the very-nearly-constant hypersonic drag coefficient. 

Where to obtain such information for the inputs is also summarized in Figure 3.  The Justus and Braun reference has atmosphere models for using this kind of analysis at a variety of places around the solar system.  The author’s spreadsheet file has separate worksheets corresponding to the scale height models and entry interface altitudes for Earth,  Mars,  Titan,  and Venus,  all from Justus and Braun.   

The author also has another spreadsheet file that does the classical 2-body orbital mechanics of elliptical orbits.  This is the best kind of source for speed at entry interface.  Technically,  evaluating slopes at the entry interface location will get you the entry angle below local horizontal,  but a default guess of 2 degrees is rather representative for spacecraft items.  Some warheads come in steeper,  but if so,  usually slower,  too,  because they are fundamentally suborbital. 

Masses and dimensions for many craft can be found on the internet.  The author has found the old Hoerner “drag bible” reference a good source for drag coefficients. 

Figure 3 – Typical Sources of Data

Ballistic coefficient β = M/(C_D A) is a measure of how well the vehicle penetrates through the air while decelerating.  If the hypersonic C_D is a constant,  then the hypersonic beta will be constant,  which is what the Allen and Eggers model assumes.  That assumption is at least approximately true all the way down to about local Mach 3 for blunt shapes. 

The dimensions and shape enable calculating a volume corresponding to the outer shape envelope.  Dividing mass-at-entry by that volume gets you an “effective density” for the craft.  Not all craft will have the same “effective density”:  manned craft will compute lower because of the interior volumes required to be open space in which the astronauts can live.  Unmanned craft will typically have higher “effective densities”,  because things can be packed as tightly as possible. 

As indicated in Figure 4,  ballistic coefficient β ~ eff. density * dimension³/dimension²  = eff. density * dimension,  for any given shape,  since volume is proportional to dimension cubed,  while area is proportional to dimension squared.  That means for the same shape and density,  ballistic coefficient scales as the cube root of mass at entry.  The same shape corresponds to the same blockage area-basis drag coefficient.

Figure 4 – How Ballistic Coefficient Varies With Mass,  Density,  and Dimensions

The author’s entry analysis spreadsheet is depicted in Figure 5 below.  This particular one is for the Earth atmosphere model,  for an Apollo coming back from the moon.

Highlighted in yellow near the top of the worksheet are 3 groups of inputs.  Of these,  the user need only worry about 2!  The leftmost group has the atmosphere model,  and there is one worksheet for each different atmosphere model.  Currently,  there are worksheets for Earth,  Mars,  Titan,  and Venus.  The atmosphere model has ρ₀ and h_scale for the exponential density variation model,  plus the entry interface altitude.  It also has the upper and lower values of altitude,  between which the scale height density approximation best matches reality.  There is an input with the name of the world the worksheet models.

The center yellow input group is the user-input entry conditions:  speed at entry interface V_atm,  and angle below horizontal Ɵ.  There is an input denoting what the mission is about. 

The vehicle model is the rightmost yellow input group near the top.  It has values for ballistic coefficient β and the effective nose radius Rn.  There is also an input for the name of the vehicle. 

The heating model constants are also given for convection and for plasma radiation.  These are not yellow,  and are not user inputs.  They are as meant,  for each worksheet.

The main calculation block starts in the left column with a list of altitudes highlighted green,  that starts at its top with the input entry interface altitude.  The user may freely adjust that list to get points denser in distribution where speeds,  gees,  and heating rates are changing rapidly.  It ends with a yellow highlighted user input altitude to find exactly the altitude that corresponds to the intended end-of entry speed.  Mach 3 on Earth is typically right at 1.0 km/s.  Mach 3 on Mars is typically close to 0.7 km/s. 

Figure 5 – Appearance of the Author’s Entry Spreadsheet Worksheets 

Typical spreadsheet results  are shown in Figure 6 just below.  These plots are generated automatically by the worksheet.  The user can see where the points need to be denser when adjusting his altitudes list.  Then when done,  he can copy these plots and paste them into a “Paint 2-D” png file.  It is recommended to read values out of the worksheet calculation block,  and annotate the resulting plots with them,  once they are in the png file.  

Figure 6 – Image of Png File Containing Annotated Worksheet Plots

The convective and radiative heating models currently embodied in the spreadsheet file’s worksheets are illustrated in Figure 7 just below.  The original Allen and Eggers model had only the convective stagnation heating model.  The author found one for plasma sheath stagnation radiation heating in the SAE Aerospace Applied Thermodynamics Manual (1969),  modified it slightly,  and incorporated it into the spreadsheet.           

The figure also has the old entry engineer’s “rule of thumb” that says the effective temperature in degrees K,  of the plasma sheath near stagnation,  is numerically equal to vehicle speed in meters/second.  This is rather crude,  being only about 10% accurate,  but it is “in the ballpark”. 

The figure also includes the author’s wild guesses for how to rescale the stagnation heating rates to other locations on the vehicle.   There are regions of attached flow that feature severe flow “scrubbing” of the surface,  and separated-wake locations that do not.  The plasma radiation heating rescales differently than the convective heating.  

For regions where flow is still attached,  the plasma sheath is still crudely as close to the surface as it is at stagnation,  implying radiation heating rates still very near stagnation,  unlike convective.  In the wake,  the plasma sheath is remote,  but still “shining upon” the surfaces,  so the author does not rescale it down as far as he does the convective. 

Figure 7 – Stagnation Heating Models Currently In the Spreadsheet,  Plus Scaling Elsewhere

Complicating Factors:  tumble-home angle vs angle-of attack for capsule shapes             

Most capsule shapes have what is called a “tumble-home angle” of the lateral walls inward.  Flow usually accelerates sub-sonically,  radially outward behind the bow shock,  to a sonic line that is usually at the very rim of the heat shield.  Flow usually separates at the rim,  just downstream of the sonic line,   leaving the lateral walls in separated wake flow,  if the capsule flies straight with no angle of attack. 

A modest angle of attack to create a lateral lift force has been used for a long time (since Gemini in the 1960’s) to better “fine-tune” the entry trajectory.  One just rolls the capsule to point that lift vector in the desired direction.   This has the effect of reducing the angle between the lateral wall and the separated flow on the side where the stagnation point is closest to the rim.  On Apollo,  this had the effect of flow staying attached to the lateral wall (with higher heating) in a localized swatch of surface,  on that side.  This sort of thing is depicted in Figure 8 just below. 

The simple entry model does not handle such subtle differences,  it just pulls the capsule straight in,  along a straight line in Cartesian coordinates,  and it only estimates stagnation heating.  The user has to allow for this possibility,  when rescaling stagnation heating rates to lateral walls where flow might actually be attached!

Figure 8 – Effects of Modest Angle of Attack Upon Heating for Capsule-Type Shapes

As an example of this angle of attack effect,  consider the data the model predicts for Apollo coming back from the moon,  in Figure 6 above.  Stagnation convective was 144 W/cm²,  and radiative was 236 W/cm²,  for a stagnation total of 380 W/cm².  Those numbers scale for attached flow locations to 48 W/cm² convective,  and 236 W/cm² radiative,  for a total of 284 W/cm².  For separated wake zones,  those same numbers rescale to 14.4 W/cm² convective,  78.7 W/cm² radiative,  for a total of only 93.1 W/cm². 

Note that the rim of the heat shield would definitely be a region of attached flow,  at total heating 284 W/cm²,  some 74.7% of that at the stagnation point!  At some angle of attack causing flow attachment for a swatch along only one lateral side,  the same high heating at something like 284 W/cm² would exist!  The rest of the lateral sides are all in separated flow,  at a heating rate only in the neighborhood of 93.1 W/cm²,  only 24.5% of stagnation.

The lesson is quite clear:  lateral sides that might see attached flow at angle of attack,  require thicker heat protection than those that do not!  That increased thickness requirement is at least similar to the thickness near the rim of the base heat shield!

Max pressure on the heat shield is important for choice of an adequate material,  as some can be crushed.  You have a mass at entry,  and a blockage area,  in order to set up your calculation of ballistic coefficient β.  The entry model spreadsheet gives you an estimate of the max deceleration gees.  Mass * max gees * gc  equals the decelerating force F acting upon the vehicle.  Max deceleration occurs high enough up,  that backside pressures on the aft surfaces are essentially zero.  So,  the average pressure on the heat shield is simply that deceleration F divided by the blockage area.  The sonic pressure near the rim is roughly half the stagnation pressure,  so the average pressure is roughly ¾ of the stagnation pressure.  Reversing that leads to P_stagn = (4/3)*P_avg,  as indicated in Figure 9.  

Figure 9 – Approximate Stagnation Pressure Estimate

Heat shield materials have definite operating limits.  Ablatives are usually rated to max heating rates per unit area,  and max pressure exposure,  as shown in Figure 10 just below.   Transpiration surfaces would likely be similarly rated,  although that technology has yet to fly (but it might soon).  Refractories are usually rated somewhat differently,  being rated directly in terms of a max service temperature,  although there is still a max pressure rating.  The user should be aware that these max rating values recommended for ablatives will vary from source to source. 

Looking at the Apollo lunar return example above,  the exposures and the ratings for its Avcoat 5026-39 heat shield compare as follows:

Item…………….exposure…….rating…….remarks

Q/A, W/cm²…..380…………….600……….OK

Max P, atm……0.56……………0.50………barely not OK,  but it worked

Figure 10 – Max Rating Values for a Few Ablatives (Values in Other Sources Vary)

The variation in ratings from source to source can be seen comparing Avcoat 5026 in Figure 10 above to “Avcoat” for Apollo and Orion in Figure 11 just below.  Note particularly the manufacturing difference between Avcoat for Orion EFT-1 versus Orion as flown in Artemis.  Artemis leaves out the reinforcing hex,  to get bonded tiles instead of hand-gunned honeycomb cells.  Most such sources leave out sufficient clarifying details!

Figure 11 – Many Ablative Applications and Rating Data (from a different source)

Ratings for some refractory ceramic materials are shown in Figure 12.  The first 3 in the figure were used on the space shuttle.  The windward tiles were colored black to increase their thermal emissivity,  where heating was larger.  The leeward tiles were white where high emissivity was not required,  but solar reflectivity was required,  for passive thermal balance control. 

These were very low density aluminosilicate materials,  whose max service temperatures were not limited by melting,  but by a solid phase change causing shrinkage and fatal embrittlement.  That last is exactly why Coleman gasoline lantern mantles were so fragile!

The ceramic blankets were more sharply temperature limited,  and were only used on leeside surfaces immersed in separated flow.

Tufroc is not a single material,  but two layers of different ceramic materials mechanically coupled together.  These are usually set up as two-part tiles bonded to the surfaces they protect.  The outer surface layer is a denser,  more thermally conductive ceramic that is rated to a higher temperature than aluminosilicates,  and also quite a bit stronger than the shuttle tile material.  The inner layer is somewhat similar to shuttle tile,  being low density,  not as strong,  and very low thermal conductivity.  It is rated to a bit-higher temperature.

Figure 12 – Some Data on Refractory Ceramic Materials

Exposed metals are possible,  but only if the heating rate is low enough to permit a survivable equilibrium temperature,  with a hot strength that is still acceptable.  This was done on Mercury and Gemini,  which returned only from low circular Earth orbit where the heating rates were far lower.  This could not be done with Apollo,  which returned from the moon at very near escape speed,  with very much higher heating rates.  It is being done again by SpaceX with its “Starship” leeside surfaces,  but only in separated flow zones,  and only from low circular Earth orbit speeds (at least so far).  See Figure 13 for materials data.

Figure 13 – Some Data on Exposed Metals as Refractory Candidates

For ablatives,  refractories (ceramics and metals),  and transpiration-cooled designs,  the heat balance concepts,  as simplified,  are shown in Figure 14 below.  These are couched in terms of heat flux format,  that being heat flow rate per unit of exposed surface area.  That matches the output data from the entry spreadsheet model. 

For the ablative scenario,  there is both ablation and re-radiation cooling available to establish equilibrium,  but no adequate way to determine how much of each!  For the refractory scenario,  there is only re-radiation cooling,  and an equilibrium temperature is easily determined iteratively.  For the transpiration scenario,  equilibrium surface temperature is constrained by coolant vaporization at an acceptable coolant pressure.  Thus,  an actual coolant flow rate is determined from that acceptable temperature.

Bear in mind that transpiration cooling has yet to actually fly in space.  It was supposed to be investigated with the old X-20 “Dyna-Soar”,  that was cancelled without ever flying.  However,  such a thing may well fly soon.  There is at least one “new space” competitor that wants to use it,  and it was seriously considered by SpaceX,  before they went with very slow-ablative tiles on their “Starship”. 

Those notions lead directly to the guidance for spreadsheet-based heat balances depicted in Figure 15 below.  These would likely be self-generated as custom spreadsheets.  This author has none to offer at this time.  The ablatives scenario must have some other constraint in order to set the point on the regression rate vs equilibrium temperature trend.  

Figure 14 – How the Heat Fluxes Balance for the 3 Scenarios

Figure 15 – Guidance Toward Setting Up Spreadsheet Heat Balances for the 3 Scenarios

Mars entry is definitely different from Earth entry,  as shown in Figure 16 just below.  These are the cross-plotted results from a study run with these tools.  The author “made-up” a small probe,  with either a conical or a blunt heat shield,  and ran it for free direct entries off an interplanetary trajectory at Mars,  plus low circular Earth orbit entries,  and entries at near escape speed.  These data were combined with results from an earlier Apollo entry study that included entry from low circular orbit and near-escape returning from the moon. 

The 2 left-side plots in the figure basically show the effect of the very thin Mars atmosphere upon end-of-hypersonics altitude,  and upon estimated stagnation pressure on the heat shield.  The surface density on Mars is numerically the same as density near 35 km altitude at Earth.  The plot of stagnation total heating vs speed at peak heating shows no reliably-discernable trend,  except that peak heating speed is higher if entry interface speed is higher.  The Mars data fall right in the middle of the Earth data,  all for comparable entry speed,  since direct entry speed at Mars is about the same as low Earth orbit entry speed.  

Figure 16 – Comparison Cross-Plots for Earth vs Mars Entries

Doing these kinds of entry studies using pencil-and-paper engineering,  assisted by modern spreadsheet software,  is actually easier than most people think.  But the engineering analyst who does this must really know what he/she is doing!  This is very heavy into high-speed compressible flow analysis,  and very high-speed heat transfer techniques! 

Plus,  in order to function,  the engineering analyst must know an awful lot about materials,  their properties,  and their limitations! 

But,  there is an undiscussed advantage if the engineering analyst can really do this pencil-and-paper engineering stuff!  He/she will have enough experience from running such numbers for many projects,  to spot bad results coming from someone else’s code.  Computers process bad inputs and bad models into bad results,  as easily as they process good inputs and good models into good results.  They all look the same,  at first glance!

References as indicated above:

#1.  H. J. Allen and A. J. Eggers,  “A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the Earth’s Atmosphere at High Supersonic Speeds”,  NACA Technical Report 1381,  44^th Annual Report of the NACA 1958,  Washington D.C. 1959. (unclassified) – this has the scale height atmosphere model and the relationship between altitude and velocity,  plus the convective stagnation heating correlation.

#2.  C. G. Justus and R. D. Braun,  “Atmospheric Environments for Entry,  Descent,  and Landing”,  MSFC-198,  June,  2007.  – this has the same Allen and Eggers entry model,  and scale height atmosphere model as Allen and Eggers,  but goes beyond just Earth.  Atmospheres for Mars,  Titan,  and Venus were obtained from here.

#3.  SAE,  “Aerospace Applied Thermodynamics Manual”,  1969.  (hardbound) – this had a simple plasma radiation heating model that was modified and added to the spreadsheet embodying the Allen and Eggers technique.

#4.  Sighard Hoerner,  “Fluid Dynamic Drag”,  self-published by the author,  1965.  – drag data for many shapes into the low hypersonic range are in this reference.

-----  

Search code DDMMYYYY format     01052026

Search keywords   aerothermo,  space program

-----   

PS:  This article actually comprises a pretty good user’s manual for my current version of the entry spreadsheet.  This spreadsheet is available from the New Mars forums as a free download,  or you can contact me directly by email.  Watch this site for two follow-up articles done by using this spreadsheet-based analysis technique. 

One will be a comparative re-entry study done for typical Mars probe heat shield shapes and an Apollo capsule shape,  all with ablative heat shields,  done at both Earth and Mars.  It will show how Mars entry is different,  with some indications as to why.

The other will be a heating distribution study for an Orion capsule doing free-entry returns from the moon.  Such will be useful for understanding what effects showed up on the Artemis-1 heat shield versus the Artemis-2 heat shield,  and the original Orion EFT-1 test flight’s heat shield.

-----   

 

Preliminary Evaluation of Artemis-2 Heat Shield

Not much credible information is yet available,  as only a handful of official NASA photos have been released.  There is one unofficial CBS news photo,  whose digitally-enhanced zoom-in,  fills-in some of the information gaps.  But I do my best with what there is!

Figure 1 depicts the expected re-entry flow and heating pattern for a lunar return by Orion,  which “flies” at angle-of-attack,  to generate a small lift force that is used to fine-tune the trajectory shape.  It does this by rolling about the wind vector,  to point the lift vector in the desired direction.  This was also done on Apollo,  and on the earlier Gemini missions.

Figure 1 – Expected Flow and Heating Pattern for an Orion Lunar Return

The Orion capsule has now flown 3 times as of this writing,  once before the Artemis program even existed!  All 3 were re-entries at near escape speed,  which is what a lunar return actually is!  That first test was called EFT-1,  launched using a Delta-IV rocket (now retired).  It had the heatshield fabricated from Avcoat,  hand-gunned into the cells of a fiberglass hex bonded to the capsule outer shell,  just like Apollo!  The other two Orion flights (so far to date) were launched with the SLS rocket,  and had cast and machined Avcoat tiles bonded to their outer shells,  unreinforced by any fiberglass hex!

Figure 2 below is a photo of what I believe to be the Orion flown on EFT-1.  I cannot vouch for the pedigree of this photo,  or that some of the char has not been deliberately removed for investigating the material beneath.  But the presence of the hex is clearly visible,  so this cannot be either the Artemis-1 or Artemis-2 Orion capsules!  If an Orion at all,  and I believe it is,  this has to be EFT-1! 

Note that a fair amount of ablation and erosion damage seems to be evident on the lateral side.  I cannot say for sure,  but this would appear to be the lateral side nearest the offset stagnation point,  where attached flow would feature heating comparable to that at the rim of the base heat shield.   The base heatshield itself is not visible in this view.

Figure 2 – Photo,  Pedigree Unknown,  of What is Thought to Be EFT-1

NASA changed the manufacture of the Orion capsule heat shield for Artemis,  retaining the Avcoat material,  but deleting the reinforcing hex,  to save time,  effort,  and money.  They built both the Artemis-1 and Artemis-2 heat shields this new way,  without waiting for the results of the uncrewed test flight that was Artemis-1 (and THAT was their fundamental mistake).  They were surprised by the unpredicted nature of the damage that Artemis-1 exhibited!  It shed both small and large chunks of char during re-entry,  leaving alarming craters in the heat shield!  Some of those are shown in Figure 3 below. 

This Artemis-1 flight was a skip re-entry with two heating episodes separated by a slight cooldown,  conducted at a very slightly shallower-than-normal angle,  compared to a straight-in re-entry.  Shallower reduces total heating a little,  but it does incur at least some risk of bouncing off the atmosphere like a skipped rock,  into an extended elliptical orbit,  whose period (of 5-10 days) exceeds the remaining crew life support duration! 

Flying the capsule at angle of attack is a way to control that skip effect.  Early in the re-entry,  you point the lift vector down,  to stop any skipping-off.  Late in the re-entry,  you point the lift vector up,  to keep the trajectory from “drooping” downward too soon.  

Figure 3 – The Official NASA Photo of the Artemis-1 Heat Shield,  After Recovery

After Artemis-1,  NASA spent nearly 2 years doing tests and calculations to “officially” convince itself that if they deleted the skip and came in slightly steeper (at higher heating),  that the alarming chunk-shedding would not occur!  This was based on the hypothesis that the pyrolysis gases produced during the second heating pulse could not percolate through the char layer easily enough,  and ended-up cracking it,  and blowing-off chunks of it,  with the obstructed gas pressure from beneath.

That hypothesis ignores the effects of fluid-scrubbing shear action in attached flow,  which would want to peel chunks off from between any cracks.  It also ignores any embrittlement and weakening of the charred material,  after cooling down some,  between the heating pulses.  Anyone who has ever dealt with the fragile mantles of a Coleman gasoline lantern,  would know exactly what embrittlement effect I am talking about,  and that it is quite real!

So,  NASA flew Artemis-2 crewed,  with the very same heat shield as Artemis-1,  just deleting the skip in favor of a slightly-steeper straight-in re-entry.  Their analyses said that would eliminate the chunk-shedding.  The real question is,  did that really work?

The few official photos released so far say that it did work.  However,  there is an unofficial photo that says “maybe not near as well as assumed”.  You judge for yourself. 

Figure 4 is an official photo taken during crew extraction from the capsule,  floating in the sea.  Bear in mind that you cannot see the capsule base heat shield at all in this view,  and you can only see the lateral side where the windows were located.  Those have to be away from the worst lateral heating,  to protect them from being destroyed by that heating.  

Figure 4 – Official NASA Photo of Artemis-2 During Crew Extraction

Figure 5 is an official NASA photo of mission commander Reid Wiseman pointing at some kind of an eroded crater in the lateral wall ablative insulation.  You can see the window behind his head,  and you can see there is no heavy charring (black) in this view.  This view is of the windows-side of the capsule,  where heating and its effects are greatly reduced by being in separated wake flow.  You can see nothing of the base heat shield at all,  or the more highly-heated opposite lateral side,  where flow was attached. 

Figure 5 – AI-Doctored Version of an Official NASA Photo of Mission Commander Reid Wiseman Pointing to Damage Spot

Figure 6 is an official NASA photo taken by a Navy diver,  of the base heat shield,  while the capsule was still in the water.   It looks rather pristine.  However,  you can see by the streaks that the stagnation point was just out of view,  top center left of photo.  The odd feature lower left is the mark or trace,  left on the heat shield from the flow about one of the tie-down pads. 

We see no craters from lost chunks of char in this view,  which supports the conclusion that NASA was right to delete the skip in the re-entry trajectory!  But we cannot see the region nearest the stagnation point,  where heating is higher,  and where fluid shear is higher.  That would be because the acceleration to sonic at the rim,  takes place over a much shorter distance.  That raises the surface shear forces that the heat shield material and its char layer “feel”. 

A suspicious person might say that only photos supporting NASA’s hypothesis have been released.  That is because we have not seen any photos of the offset stagnation region on the base heat shield,  or of the lateral wall on the side adjacent to that stagnation region (opposite the windows),  where heating is almost as high as on the base heat shield.  Know that eventually,  all the photos must be released,  in the report on their heat shield investigation.  That report must be publicly released!  But it may be a year before we see it!

Figure 6 --  Official NASA Photo of Artemis-2 Heat Shield Taken by a Navy Diver

As I said above,  there is an unofficial photo now circulating,  that was taken by CBS News seconds before splashdown,  while the capsule was still hanging from its parachutes.  It was distant and somewhat blurry,  but it does show some sort of white mark near the heat shield rim. 

That white mark attracted a lot of attention,  and led to the digital enhancement of that photo,  and a digital zoom-in to examine that white mark more closely.  But that enhanced photo does indeed also show the near-stagnation region of the base heat shield,  and the lateral wall away from the windows,  where flow was attached,  and the heating much higher.  That original blurry CBS News photo is Figure 7 below,  and the zoomed-in enhancement is Figure 8 below. 

You cannot see very much in Figure 7,  but you can in Figure 8!  The white mark was left by the melting and destruction of one of the tie-down pads.  This is something NASA says was expected,  although it did not occur on Artemis-1,  as you can see in Figure 3 above.  It does seem to have left an alarmingly-deep cavity eroded into the rim of the base heat shield!  Expected or not,  that cavity would appear to be of a depth comparable the thickness of the heat shield.  And I do find that alarming!

Figure 7 – Unofficial,  Blurry Photo Taken by CBS News Seconds Before Splashdown

What nobody has been talking about are the other things I see in Figure 8.  I have circled four places that I believe show where chunks of char were shed from the base heat shield.  These are smaller chunk-shed craters,  to be sure,  and scaling up from the limited view,  not anywhere near as numerous as those seen on Artemis-1!  Yet they are there,  and the revised non-skip re-entry so very clearly did not entirely stop them from occurring!  Which simply says that something else was going on with that char chunk shedding,  besides pyrolysis gas percolation through the char! 

There’s one other circled spot,  and two arrows pointing to large locations,  on the base heat shield in Figure 8,   where I cannot tell what happened,  but I can see that some sort of damage is clearly there.  It will take better photos than this to evaluate those damages. 

What I see on the lateral side adjacent to the stagnation point are one unidentifiable small dark spot,  and two whitish bright spots of considerable size.  The two bright spots are clearly places where all the char was lost,  exposing bare metal to full heating!  And that metal looks distorted by that heating!  Very alarming indeed!

Those bright spots are not windows,  those are the bare metal of the outer capsule shell,  to which the Avcoat tiles were bonded!  Simple thermal insulation separates it from the inner metal shell,  which is the capsule crew cabin pressure vessel. 

Figure 8 – Digital Enhancement and Zoom-In of Unofficial CBS News Photo

There would seem to be four things going on here that affect the damages seen,  not just the one thing that NASA hypothesized.  They are:

#1. Pyrolysis gas percolating out against permeability resistance,  wanting to blow chunks off. (That is the NASA hypothesis.)

#2. Fluid surface shear forces wanting to peel chunks out from between cracks in the surface. 

#3. Char layer shrinking,  cracking,  and embrittlement upon cooldown,  between the heating pulses of a skip-type re-entry.  (Like the fragility of a gasoline lantern mantle.)

#4. The presence of the reinforcing hex actually ties the char layer tighter to the pyrolysis and virgin layers beneath,  and it also acts to limit the spread of cracks in the surface.

My conclusions about the rational things to do,   depend upon what response NASA takes to this Artemis-1 and -2 outcome:

#1. If NASA ever wants to resume flying skip-type re-entries,  then put the reinforcing hex back into the Avcoat!  Period!  There actually is a way to do that,  without hand-gunning Avcoat into every hex cell like Apollo!

#2. Non-skip re-entries have higher peak heating.  The Avcoat tile thickness is insufficient at the attached-flow locations:  near the tie-down pads,  and on the lateral wall opposite the windows.  NASA must thicken it at those locations!

Avcoat tiles with reinforcing hex,  but without hand-gunning:

Avcoat is an epoxy-novolac polymer loaded with solids.  Those solids include some small amount of carbon fibers,  and a lot of tiny micro-balloons made of phenolic resin.  Higher micro-balloon content lowers density,  raises ablation rate,  and increases the porosity and permeability of the char layer (also decreasing its strength).  It also greatly increases the apparent uncured mix “viscosity”,  which is already extremely thixotropic (almost a solid). 

The Avcoat mixture is so thick,  that it is almost “crumbly” coming out of an air-powered caulking gun,  whose nozzle matches the size and shape of the hex cells.  The hex is a fiberglass cloth with a phenolic resin matrix.  The glass softens at a higher temperature (~ 900 F) than that at which the polymer starts to pyrolyze (~ 300 F),  which is why it is an effective char layer stiffener and retention aid.  The carbon fibers add a bit to those reinforcing effects.  Apollo and Orion EFT-1 used hex panels bonded to the outer shell,  into which cells the Avcoat was hand-gunned.  This was extremely labor-intensive!

For the Artemis-1 and Artemis-2 Orion capsules,  the Avcoat mixture was cast into blocks,  from which tiles were machined.  There was no reinforcing hex!  The bonds and gap fillers of those tiles are not in question here,  those performed just fine!  The lack of reinforcing hex is the question!

You can make tiles with hex in them,  but only if you can stop being bound by “either/or thinking”.  That would be either doing it the Apollo/Orion EFT-1 way,  or doing it the Artemis-1/-2 way.  However,  there is a third path!  Read on:

Put a chunk of the hex the size of the cast block you want to make,  into some tooling on the outlet of a plastics extrusion press.  Load your Avcoat mixture into the press,  and use that press to force it through all the cells in the hex,  all at once!  Remove the loaded-hex tooling from the press,  put the bottom and top on that tooling,  and cure that block of Avcoat that now contains the reinforcing hex!  Then,  machine your tiles from those blocks,  and bond them to the Orion outer shell,  the same way as in Artemis-1 and -2!  Tiles that are hex-reinforced,  but with NO hand-gunning!

I gave this idea to the NASA heat protection group in Houston well over a year ago,  as of this writing,  and again directly to the new NASA Administrator Jared Isaacman only a few weeks ago.  I was able to confirm (1) that the heat protection group got it,  and (2) that they thought I was right.  Nothing confirms that Isaacman ever saw my letter to him. 

But I never heard another word out of NASA about this alternative,  and NASA has not yet done anything like it.  So,  I must conclude:

Money and schedule clearly still outweigh crew’s lives for the decision-making NASA management levels.  And “not invented here” is still quite strong at NASA,  as well as its contractors.  Those two things are my real ongoing reservations about NASA!  And they have been,  ever since the first of two lost Space Shuttle crews!  Crews lost precisely because of those very same two management culture flaws!

-----  

Search code DDMMYYYY format:     28042026

Search keywords:       aerothermo,  space program

-----   

Closely-Related Postings:

“About the Artemis-2 Mission” posted 31 March 2026

The updates to that article have these same photos in them,  just not as explicitly annotated as here.  And I have since added how the flow and heating patterns vary.  In near-escape Earth entry with a blunt heat shield,  plasma radiation heating is larger than convective heating,  and it decreases with distance from stagnation less rapidly.

Search code DDMMYYYY format      31032026

Search keywords         aerothermo,  launch,  radiation,  space program 

“Ramjet Data Re:  Heat Shields” posted 1 March 2026

Shows how my old experiences with ablatives in solid rocket motors,  and especially ramjet combustors,  have strong overlap with the re-entry phenomena involved here.  Char retention by the layers below,  is a crucial key,  in both venues! 

Search code DDMMYYYY format      01032026

Search keywords         aerothermo,  ramjet,  space program

-----  

 

 

Trump the Incompetent Liar

 I originally wrote this as a column and submitted it to the Waco "Trib" newspaper almost a week ago.  They have so far not chosen to use it.  So I have posted it here.  

-----   

Trump lied about the economy.  He actually inherited a recovering economy with low inflation,  and then damaged it severely with his tariffs and trade wars.  Further,  you are now facing simultaneous global recession and out-of-control inflation,  because of the closure of the Strait of Hormuz. 

Trump lied about tax cuts.  You got a small one (if any at all),  while his billionaire and corporate cronies got billions!  So,  the national debt is going completely out of control,  especially with high inflation driving up interest rates.  The interest on the debt now dominates the annual “budget”.  Which proper budgets Trump’s enablers in Congress repeatedly refuse to pass.

Trump lied about immigrants!  Most of them,  including most that are here illegally,  are more law-abiding and honestly tax-paying,  than the bulk of you readers out there!   So say the crime statistics!   And those same people are integral to our economy.  They harvest your food,  build your homes,  mow your lawns,  and serve as housekeepers.  And more!

Yet Trump’s lies about how evil all immigrants are,  have been the “justification” to weaponize ICE and CBP into something to abuse and “disappear” people that he does not like,  while trying to intimidate the rest of us into submission!  That is the very definition of a dictatorship’s secret police force!  And while you may have voted for Trump,  you did NOT vote for that!

Trump has lied about everything since the start of his Iran war,  including especially the “why”,  and the “why now”.  The experts dispute quite strongly that the Iranians were anywhere near ready to build a nuclear bomb.  Trump was frustrated because he couldn’t get a deal about the enriched uranium (after he himself scrapped the one we had before).  Then Netanyahu walked up wanting to start a war,  because of Iran’s proxy Hezbollah. 

Trump lied about how thoroughly we destroyed Iranian military capabilities.  Those lies are painfully evident.  While reduced in numbers,  Iran can still strike targets all over the region,  and ships in the Strait of Hormuz!  Which was open until Trump started his Iran war!  And now,  we have mostly run out of interceptors,  so those fewer Iranian weapons get through!

Trump has been lying about the progress of the peace negotiations with Iran!  Evidently,  he did not even know who he was really dealing with!  Or why the things he thinks he obtains during these talks,  repeatedly get quickly taken away. 

Those Iranian officials at the table in Islamabad are NOT the de-facto rulers of Iran now,  the Iranian Revolutionary Guard Corps (IRGC) is!  The IRGC is a huge,  well-armed,  terrorist army of violent extremists,  recruited specifically because they are violent extremists!  They would die to the last man,  before stopping the use of the other proxy terrorist armies (Hamas,  Hezbollah,  the Houthis,  and many more),  before giving up their enriched uranium for their terrorist bomb,  or before opening that Strait!

Trump simply does not understand non-transactional people like that,  not at all!  And he (and Netanyahu) actually made the IRGC the de-facto rulers of Iran,  by killing off most of the clerics and many of the regular government officials,  in the first few days of this war! 

Further,  by bombing civilian areas,  Trump has essentially united the people of Iran against us,  instead of helping them to overthrow their regime.  Regime change is actually the only way to get what we’d like to have from Iran!  It is now out of reach!

It appears to me that Trump started this Iran war because he could not get the then-ruling clerics in Iran to deal transactionally with him.  Just like Maduro in Venezuela.  And soon Cuba,  if and when Trump gets clear of his Iran debacle! 

In order to get clear,  Trump must now commit war crimes by deliberately bombing civilians and their infrastructure,  all across the country!  All the aboveground military targets actually have been destroyed,  although clearly many remain,  deep underground! 

You who voted for him,  you did NOT vote for American war crimes!  America’s name is now mud around the world,  because of the stupid things Trump has done,  including to our allies!  They no longer trust us at all,  and justifiably so,  nor do they even want to trade with us,  anymore!  And one of them,  Ukraine,  has developed a cheap,  mass-produced anti-drone drone,  that we do not have!  In 20-20 hindsight,  Trump’s mistreatment of Zelenskyy,  and starving Ukraine of assistance against Putin’s invasion,  looks very stupid indeed!

There is absolutely no excuse for that level of lying,  that much chaos,  and that egregious level of incompetence,  which we have seen out of Trump and nearly every appointed figure in his entire administration!  And all of you readers bloody well know that what I said is true!

Now,  readers,  please go out and do your job as proper citizens in November!  Vote all of Trump’s enablers out,  so that we can get rid of all this chaos and incompetence before it destroys us,  and get back to being a responsible world power.  With a decent economy.  But without almost-a-king!

-----   

below is a photo of a Ukrainian-developed drone interceptor,  plus some more comments

-----   

A Ukrainian Shahed interceptor drone. Photo credits: Bohdan Myroshnychenko

This drone is actually cheaper to build than the Iranian "Shahed" drones that it can destroy.  The Ukrainians developed this on their own,  faced with chronic mass "Shahed" drone attacks from Russia.   The Iranians have been supplying Russia with those "Shahed" drones for some time now.

Those who know me,  know that I consider what Trump has done to our alliances and our allies,  to be treason of the aid and comfort type.  The most egregious case has been trying to ram Russia's terms down Ukraine's throat,  in his so-called "peace process" negotiations.  But every alienated ally is another count of that treason.  

Those who know me also know that I consider Trump's egregious and infamous executive over-reach to be really the attempted establishment of a Trump dictatorship over us.  He is doing that so that he won't have to leave office until he dies.  That would leave us with a whole series of Trump-wannabees as dictators over us.  I estimate that imposition to be about 70% complete now.

The weaponization of ICE and CBP into a secret police force that "disappears" some of us,  and intimidates the rest of us,  is a piece of that dictatorship.  As are the "detention centers" which are really concentration camps for those waiting to be "disappeared".  As is the identification and vilification of some group,  as the enemy to be "disappeared" (in Trump's case,  immigrants).  All such are features of all dictatorships.  

Do not listen to the words,  from anybody!  Look only at the events and actions!  They verify everything that I have claimed.  

-----   

search code DDMMYYYY format:    24032026

search keywords     bad government, idiocy in politics, treason

-----   

Elliptic Departure and Arrival

Update 4-22-2026:  This post is fundamentally about using extended elliptical orbits and "space tugs" to reduce the departure (and arrival) velocity requirements on lunar and interplanetary craft of any type.  That technique provides significant benefit even if you throw away the tug stage,  as Artemis-2 just demonstrated. 

However,   if you combine it with a reusable "space tug" stage,  plus a space station in low circular where that space tug stage can be based,  the benefits greatly magnify.   Such a station needs to do both lunar or interplanetary craft assembly,  and the filling or refilling of such craft and the tug with propellants.  It does not require continuous manning,  unless other tasks needing that,  are also done there.  That space station base is also covered in the article.

The critical enabling factor will be propellant transfer from tank to tank in zero-gee!  At a space station,  you do NOT want to do such transfers by spinning large structures,  or with orbit-altering ullage thrust!  We can already do that task for storable propellants with bladder-expulsion tanks.  But there has not been such a solution for cryogenic tanks.  

Until now.  

I came up with a scheme that would work quite well.  The article covers it,  as well. 

This article thus gives you the path to an affordable,  but very capable,  space program!  There is now no excuse not to do this. 

-----   

original posted article follows

-----   

Artemis-2 used an elliptic departure orbit to greatly reduce the trans-lunar injection departure burn required of its Orion service module.  It used the SLS core stage and SRB’s to enter a surface-grazing ellipse with a very-low apogee,  and then used its ICPS second stage to raise the perigee of that ellipse out of the atmosphere with a small ICPS burn at apogee.  Upon reaching the new perigee,  it made another substantial ICPS burn to raise the apogee into a very extended departure ellipse.  It demonstrated what I propose here!

The ICPS had just enough propellant to be disposed-of safely,  after the Orion capsule and service module separated,  after getting onto the extended ellipse.  The Orion and its service module made one circuit about the extended ellipse,  preparing for the lunar trip.  Upon reaching perigee,  a modest service module burn put it onto the lunar trajectory,  with propellant reserves for course corrections both ways.  This lunar trajectory was a free-return loop around the moon,  with a direct re-entry back at Earth.  See Figure 1. 

Figure 1 – How Artemis-2 Used Elliptic Departure to Reduce the Lunar Departure Burn

Changes for a Permanent Capability

The perigee Artemis used was still too low at about 185 km altitude,  too close to the entry interface altitude of 140 km,  to be a permanent orbit.  A more permanent low circular orbit would about 300 km or more,  as shown in Figure 2.  An extended elliptical departure orbit was selected by this author that had more than 100:1 stronger pull of Earth gravity at apogee than lunar gravity,  plus an integer ratio of its period to the period of the low circular orbit!  Such would be stable for multiple circuits,  and the integer period ratio ensures that anything left behind in low circular will be there just as you arrive back at that perigee!  That makes rendezvous and docking much easier and faster. 

Figure 2 – Proposed Elliptic Departure/Arrival Orbit With Basing in Low Circular

It would be unattractive to base directly in the extended orbit,  when its perigee speed is very nearly Earth escape speed.  The most demanding portion of the ascent is surface-to-orbit,  for which low circular is demanding enough,  just reaching circular orbit speed with any significant payload.   It takes a much bigger launch vehicle to reach near escape speed with that same payload!  That’s more expensive,  and may require dedicated designs!

Instead,  one bases in the easier-to-reach low circular,  and uses a convenient stage or vehicle as a tug,  to take its payload craft from there to the departure ellipse.  From there,  the departure burn demanded of the payload craft is quite small,  especially for lunar missions.  The “tug” for Artemis-2 was the ICPS second stage of the SLS launch vehicle.  Using a tug to get onto a departure ellipse is still a substantial reduction of the departure burn,  even for faster-than-Hohmann Mars missions,  as the figure shows.

As we already now know,  reusability dramatically lowers mission costs!  To accomplish that as a tug-assisted elliptic departure,  we need a reusable stage or vehicle to use as our tug,  and we need to base it in low circular orbit.  Such is more easily reached for re-supply from the surface.  The tug can stay on the departure ellipse,  after releasing the payload craft to make its modest departure burn.  This tug then returns around that ellipse,  and burns unladen (for low propellant expenditure!) to return to low circular, where it is based.

Basing at a Space Station

That base ought to be an appropriate space station located in the low circular orbit at low inclination (to reduce plane change requirements for lunar or interplanetary missions).  What we need of that station is twofold:  (1) the means to assemble mission craft from docked-together components,  and (2) the means by which to fill (and refill) such craft,  and the reusable tug,  with appropriate propellants. 

We already know that we need manipulator arms and a framework to support them,  from the space shuttle and ISS experiences.  We will need the means by which to transfer both room temperature storable propellants,  and cryogenic propellants,  from tank to tank in zero-gee,  without spinning big structures or using orbit-changing ullage thrust!

Given that we can accomplish those things,  the advantages are enormous,  as detailed in Figure 3.  The notation V_near is also known as c₃,  the speed with respect to Earth needed at end of burn,  close to Earth,  to accomplish the lunar or interplanetary mission.

Figure 3 – Reusable Tug From Low Circular Greatly Reduces Final Departure Burns

That means that our space station is a frame to which multiple manipulator arms are affixed,  with arm operator cabins,  mounting and holding fixtures,  plus the support equipment for such crewed activities.  It must also have a multiplicity of appropriate propellant storage tanks,  the plumbing for propellant transport point-to-point on the station,  and docking facilities for the “supply tanker” transports bringing propellant supplies to the station.

The solution to the transfer of storable propellants in zero-gee has long been known:  bladder expulsion using gas pressure,  to squeeze the bladder within the tank walls.  There are no well-known cryogenic propellant solutions,  other than large structure spin or ullage thrust,  since there are no polymers with the enormous strain capability required (over 100% elongation),  for bladder service at cryogenic temperatures!  Neither structure-spin nor ullage thrust would be useful at a space station,  for any number of reasons.

There is,  however,  a not-well-known solution involving spin,  but only spinning the propellant inside stationary tanks!  One does this with vanes,  driven by electric motors.  If one does this by spinning half the propellant one way,  and half the other way,  then all the spin reactions and gyroscopic forces sum to zero at the tank mountings!  See Figure 4.  

Figure 4 – Proposed Cryogenic Delivery Tank Using Spin of Only the Propellant

Note also that the most practical tank design,  for (only) the payload of cryogenic propellants delivered to the station,  would be that same vane tank approach!  Storables could use bladder expulsion,  same as the main propulsion tanks.  Or they could use vane tanks,  but the bladder expulsion approach is both lighter weight and long-proven.

Now the form of the space station becomes clearer:  a long truss space frame,  along a portion of which are disposed a number of storable bladder tanks and cryogenic vane tanks,  and along another portion of which there are assembly arms,  holding fixtures,  and arm operator cabins.  Plumbing and power lines get routed within the frame.  You put the crew quarters and re-boost propulsion at one end,  and leave the other open,  for tanker vehicle docking,  and for future propellant capacity growth.  See Figure 5.  

Figure 5 – Proposed Refill and Assembly Space Station Concept

Benefits of This Approach Going to the Moon

For 1-way lunar landings,  your craft only needs 0.1 km/s to depart from the ellipse where the tug took it.  It needs around 0.9 to 1.1 km/s dV to enter low lunar orbit at some convenient inclination.  From there,  about 1.7-1.9 km/s dV will land it.  That’s a total of likely-under-3 km/s dV required of your 1-way lunar landing mission craft! 

At 330 s Isp for the lunar lander using storables,  that is about a single-stage mass ratio of 2.53,  or a propellant mass fraction of 60%.  If the lander has rough field capability,  its inert mass fraction might be near 15% at most.  Which leaves a payload fraction near at least 25%,  with storable propellants!  This is how you send cargo 1-way to build a base there!

Astonishing!

And it’s only about 6 km/s to return to the extended elliptic Earth orbit,  all the way from the surface of the moon,  unrefilled!   How to return such a vehicle unladen of payload,  but as a single stage,  is a topic for another time.  But it can be done,  even with storables!

What About Mars?

For a 2-way Mars orbit-to-orbit transport using low Mars orbit,  you only need around 1.5 km/s to depart the ellipse,  and about 2 km/s to enter low Mars orbit, 1-way.  That’s only 3.5 km/s required 1-way,  and so only 7 km/s required for the complete 2-way trip,  all unrefilled in a single stage!  Which would amount to a reusable chemically-powered orbit-to-orbit transport,  even with zero infrastructure at Mars!  That’s a single-stage mass ratio of about 8.7 at only 330 s Isp for storables,  or a propellant fraction of 89%.  If the inert fraction for a vacuum-only ship is 5%,  the payload fraction could be 6%!  And with only storable propellants in a single stage!

Astonishing!

To deliver payloads 1-way to Mars for direct entry and landing,  you are looking at about 1.5 km/s to depart,  plus some modest course corrections,  and the final landing burn of 1 to 1.5 km/s.  That’s at most 3 km/s dV demanded of the craft.  At 330 s Isp with storables,  that’s a mass ratio of only 2.53,  for a propellant fraction of about 60%.  If the lander,  which must survive entry as well as be configured for rough field landing,  has an inert fraction of 20%,  that is still near 20% payload fraction,  even with only storable propellants!  That could well be how to send cargo 1-way to build a base there!

Even more astonishing! 

See also Table 1.

What About Practical Tug Designs?

Now,  what do we need of the tug?  Assuming it might have to travel the extended ellipse for 2,  maybe 3,  circuits,  that’s about 10-12 days’ time in space.  If it uses cryogenic propellants for their high performance,  it needs a useful “stage life” without serious evaporative loss, of only some 10 or 12 days!  That rules out common bulkheads between LOX and LH2 tanks,  and it rules out bare single-wall tank shells exposed to sunlight in space!  But it does NOT rule out using LOX-LH2 for its highest performance!  You just need separate LOX and LH2 tanks,  which must be well-insulated externally,  plus with a shiny foil outer covering to shade them from sunlight heating.  That’s probably closer to a stage inert fraction of 10% than the usual vacuum stage inert of 5%.

We do not need months or years of “stage life”,  only a couple of weeks,  or so!  Which means we really do not need the weight and power-required penalties of cryocooler equipment!  We just need a minimal tank redesign from what is otherwise basically Centaur stage technology.  And later on,  we might need to scale it up for larger-mass mission craft!

See Figure 6.  

Figure 6 – Probable Tug Tank Construction for LOX-LH2 Propellants and 2-Week “Stage Life”

What About Arrival Versus Departure?

The dV requirements for arrival are almost exactly the same as departure,  but the timing requirements are different!  It only takes several seconds for the tug to undock from its payload craft and move several meters away.  That means for departures,  the tug can feasibly fire for reaching ellipse perigee speed,  undock,  and let the craft fire for its departure,  all in the one pass!  The tug then makes 1 circuit about the ellipse before burning at next perigee unladen, to get back into low circular,  going back to the station. 

Or,  if mission preparation time is needed by the payload craft,  both tug and craft can make one circuit about the departure ellipse,  with the craft departing,  and the tug burning to return to circular,  at the next perigee.

Arrival is different:  the tug cannot “be there” just as the craft arrives and burns into the ellipse perigee!  The craft needs to make a circuit about the ellipse,  before the tug can (1) rendezvous with it at its next ellipse perigee,  and (2) the tug must then burn to get onto the ellipse with the craft.  It then takes significant time to actually get docked together.  So the docked pair must make a second payload craft circuit about the ellipse,  before the tug can burn,  to put them into low circular,  and take them right to the station. 

And the laden vs unladen weight statements are more beneficial for departure,  and not as beneficial for arrivals.  You’d like the bigger propellant burn to be the first,  but that cannot happen when using a tug to assist arrival.  Arrival retrieved craft sizes must then inherently be smaller,  for a given tug design.  See Figure 7 for the estimated departure and arrival data for a tug sized to put a 50 metric ton craft onto the departure ellipse defined above,  as done with simple linked rocket equation calculations in a spreadsheet. 

Figure 7 --  Typical Tug Rough-Sizing and Performance by Spreadsheet

About Resupply Tankers

The resupply tanker vehicles sent to the station from the surface could be the upper stages or payload items of almost any existing or planned launch vehicle!  If sending up a storable propellant,  the payload can be a simple bladder-expulsion tank,  quite separate from the vehicle’s main propulsion propellant.  If sending up a cryogenic propellant,  the payload needs to be a vane tank,  also quite separate from the main propulsion propellant. 

Done that way,  all the tanker vehicle need do with the space station is dock,  and then hook up to the station plumbing,  to deliver its payload.  It goes without saying that reusable tanker vehicles would be much preferred for the long term.

The main caveat for tanker vehicles is that you do not want to use the same payload propellant tank (of either type) for shipping different propellants!  Propellants do get into the inherent slight porosity,  of even metal tank walls!  You need to be able to swap out the payload propellant tank,  if the same vehicle upper stage gets re-used to deliver other propellant species!  Period!  That nevertheless could be something quite convenient,  whether the payload tank is a vane tank,  or a bladder-expulsion tank!  See Figure 8. 

Figure 8 – Plausible Tanker Supply Vehicle Configurations

Prior Related Postings

As you can see from the list below,  I have been thinking about the various aspects of,  and problems feeding requirements into,  this tug-assisted elliptic departure space station scenario,  for some time.  The dates are shown in MM-DD-YYYY format.  There is an archive search tool on the left side of this page.  All that you need in order to use it are the year,  month,  and title.  I suggest that you jot down the ones you would like to see.  Click on the year,  then the month,  then the title if there were multiple postings that month.

This site also has a keyword search option,  in the sense that if you select a keyword,  then you see only those postings that are labeled with that keyword.  The current list of keywords is:  aerothermo,  airplanes,  asteroid defense,  bad computers,  bad government,  bad manners,  cactus-killing,  climate change,  current events,  education,  ethanol,  forensics,  fossil fuel,  fun stuff,  Gulf oil disaster,  guns,  health care reform,  idiocy in politics,  IR,  launch,  Mars,  Mideast threats,  North Korean rocket test,  nuclear crisis,  old cars,  pulsejet,  radiation,  ramjet,  space program,  spacesuit,  towed decoys,  trains,  treason;  the three highlighted are the ones most applicable to this article’s topic.

Only some of the more recent postings have received a search code for direct access.  That search code is something I have to give you.  It is the posting date in DDMMYYYY format.

Date............Keywords...............................Title

11-11-2025..space program.......................Where Should the New Space Stations Be Located?

10-16-2025..space program.......................Going Back to the Moon

7-26-2025....space program.......................Tank Design for Easy Cryogenic Transfers In Weightlessness

5-01-2025....space program.......................Vehicle Assembly and Refueling Facility in LEO

1-25-2025....space program.......................Initial Study for Tug Missions LEO to LLO

1-2-2025......launch, space program...........SpaceX’s ‘Starship’ As a Space Tug

12-1-2024....space program........................Tug-Assisted Arrivals and Departures

10-1-2024....space program........................Elliptic Capture

12-9-23........launch, space program............Overall Study Results: Propellant From Moon

5-1-22..........Mars, space program..............Investigation: “Big Ship” Propellant From Moon vs From Earth

4-2-22..........Mars, space program..............Earth-Mars Orbit-to-Orbit Transport Propulsion Studies

2-1-22..........space program........................A Concept for an On-Orbit Propellant Depot

8-18-21........launch, space program...........Propellant Ullage Problem and Solutions

3-23-21........space program........................Third Spacex Tanker Study

3-21-21........space program........................Second Spacex Tanker Study

3-17-21........space program........................Spacex Tanker Investigation

7-3-20..........launch, space program............Cis-Lunar Orbits and Requirements

11-21-19......Mars, space program...............Interplanetary Trajectories and Requirements

2-11-14........Mars, space prgm,  spacesuit..On-Orbit Repair and Assembly Facility

10-2-13........space program.........................Budget Moon Missions

8-2-12..........Mars, space program...............Velocity Requirements for Mars Orbit-Orbit Missions

8-2-11..........space program..........................End of an Era Need Not Be End of a Capability

-----  

Search code (DDMMYYYY)                 14042026

Search keywords                                       space program

-----