A Closer Look At Nuclear Thermal

 This article takes a closer,  more nuanced look at nuclear thermal propulsion for large colonization ships.  It still assumes fairly large dead-head payloads,  but only carried on the outbound voyage!  Propellant is sized to make the outbound and return voyages in one stage (no stage-off or jettisoning of anything along either way,  just unload of the dead-head payload at destination).  The journey baseline is low Earth orbit to low Mars orbit,  and back.

How the ships or the payload get to low Earth orbit is unaddressed.  How the payload gets delivered to Mars’s surface from low Mars orbit is unaddressed.  How the ships are refueled and reloaded in low Earth orbit is unaddressed.  What is addressed here,  that is unlike the earlier simpler study,  are the separate inert weights associated with the payload section,  the propellant tankage section,  and the engine-with-its-associated-subsystems.  The minimum vehicle acceleration requirement is increased to 0.5 gee,  except for one system deemed adequate at 0.33 gee.

The previous closely-related article was “Colonization Ship Study”,  dated 9-9-19.  It examined the simpler-to-analyze case of carrying the dead-head payload both ways (outbound and return),  so that there was one mass ratio and one delta-vee (dV) to cover the round trip.  That scope was multiple fundamentally-different forms of propulsion:  nuclear explosion drive (or “pulse propulsion”) as it was envisioned in the late 1950’s,  nuclear thermal propulsion (as a version of the solid core NERVA for which engine prototypes were tested),  Hall effect ion propulsion based off of plentiful,  cheap,  and solid-phase-but-sublimable iodine,  LOX-LH2 chemical rockets,  and storable-propellant rockets. 

Scope here is only nuclear thermal rocket propulsion,  but with the highly-variable tested or envisioned characteristics of six different design approaches.  It is these six that are compared in terms of the ratio of initial ignition to dead-head payload weight,  using the same maximum-attractive criterion of 5 as in the earlier study.  These six approaches and their relative states of technological readiness are:

(1) as-tested NERVA solid core, 

(2) the best-anticipated solid-core NERVA derivatives that never got built or tested, 

(3) the particle bed solid core reactor engine (one version of which was “Timberwind”,  which got some exploratory testing revealing unresolved problems,  but never reached the engine prototype stage), 

(4) the so-called “nuclear light bulb” gas core concept (some insufficient feasibility tests), 

(5) the open-cycle gas core concept restricted to regenerative cooling,  meaning no radiator required (some insufficient feasibility testing),  and

(6) the open-cycle gas core concept with a large,  heavy external waste heat radiator (some insufficient feasibility testing).

To accomplish this investigation,  I added an additional worksheet to the colony ships.xlsx spreadsheet file that I used for the earlier study.  Unlike the previous study,  there are no closed-form ways to get from dead-head payload to a vehicle weight statement.  The calculation uses iterative convergence of the propellant tank inert weight,  and iterative convergence of engine thrust sizing in terms of the resulting vehicle acceleration gee capability.

For this investigation,  the payload section is presumed to be some sort of enclosed hull,  with adequate insulation,  radiation shielding,  and micrometeor protection for a crew built into it,  in some unspecified way.  The ratio of dead-head payload mass (contained inside) to the loaded payload section mass is a fraction denoted as fpay.  The dead-head payload size drives everything in the end,  as all results are directly proportional to the dead-head payload input.  For this investigation,  dead-head payload was arbitrarily set at 100 metric tons,  and fpay = 0.8,  the same for all six engine types.  Thus:

               Loaded payload section mass = dead-head payload mass/fpay

               Payload section inert mass = loaded payload section mass – dead-head payload mass

The propellant tank section contains the common propellant for all nuclear thermal engine approaches:  liquid hydrogen (LH2).  This is a harsh cryogen,  requiring solar heating control,  significant insulation,  and some sort of cryocooler to control evaporation.  This is simply going to be heavier than the lightest-possible single-wall bare tank.  The ratio of propellant mass to loaded tank mass is the fraction ftank.  The single value ftank = 0.95 was used for all six engine types.  Thus:

               Loaded tank section mass = propellant mass/ftank

               Tank inert mass = loaded tank mass – propellant mass

               Propellant mass must be the sum for two burns at differing dead-head payload

               One starts with a guess for tank inert,  and iteratively converges it to the result

The engine “section” is the nuclear thermal rocket engine (or engines,  for redundancy),  complete with turbopumps and control equipment,  a radiation shadow shield for the crew up forward,  plus any waste heat radiator that may be required (if regenerative cooling alone cannot do the job).  This radiator (if present) and the core-plus-engine hardware lead to a characteristic engine thrust/weight ratio T/We,  which is dimensionless under the definition that both thrust T and engine weight We (on Earth) are measured in force units.  This ratio is different for each engine type,  as is the resulting specific impulse.  The values I used follow:

               Type                     Isp, s      T/We               development status

               NERVA                 725        3.6               as-tested in engine prototypes

               derNERVA           1000      5               derivative-of-NERVA,  estimated on paper

               PBR                       1000      7               particle-bed reactor,  based on “Timberwind”

               Nuc.lt.blb            1300      10               “nuclear light bulb” gas core concept,  some feasibility

               Open GCR           2500      20               open-cycle gas core concept limited to regenerative cooling

               GCR+rad              6000      0.5               open-cycle GCR with heavy waste heat radiator,  concept

For this kind of data,  the main results used to size the vehicle are the exhaust velocity Vex (km/s),  and the engine system inert mass (metric tons).  These are:

               Vex, km/s = (Isp, s)*9.8067/1000

               Engine system inert mass, metric tons = thrust level, KN/(9.8067 * T/We)

For the remaining vehicle characteristics,  all the concepts except “GCR+rad” were required to size thrust level such that the vehicle acceleration at the initial ignition mass was at or just above 0.5 gee.  This corresponds to about a 15 minute Earth departure burn,  definitely short enough to qualify as “impulse”,  and not have the orbital dV be factored-up for gravity loss to be mass ratio-effective. 

With the data I used,  the GCR+rad system could not reach half a gee,  but converged fairly well at 0.33 gee.  This is less than a 30 minute burn,  still short enough to consider as “impulsive” for Earth departure.   

Max gee at final burnout weight upon Earth return should be under about 5,  but this proved not to be a problem.

It’s a two-level iteration:  first set a thrust level,  then converge your guess for propellant tank inert weight with the final result of the calculation for tank inert weight.  Then check and adjust your thrust level for the right Earth departure gee level.  Then converge the tank inert weight again.  Repeat the process as needed to get however-close a convergence you deem tolerable (0.1-0.01 ton range).

The orbital dV’s that are required are those for getting from low orbit onto a min-energy Hohmann transfer ellipse.  The values used are worst-cases that do not go together;  the difference is a nice little “kitty” to cover midcourse corrections.  Earth departure = Earth arrival = 3.84 km/s.  Mars arrival = Mars departure = 1.83 km/s.  These sum to 5.67 km/s outbound in a heavier ship carrying payload,  and 5.67 km/s return in a lighter ship with no payload and already having burned off some propellant on the outbound voyage. 

Factored for losses,  these dV figures become the mass ratio-effective dV’s for design purposes.  Those and the Vex for each engine type give you the mass ratio MR for each engine type,  one for outbound,  the other for return.

               MR = exp(sum dV/Vex)  with both velocities in km/s,  and the sum dV for outbound or return

You start the calculation with the return voyage by summing up the inerts (payload section inert + tank inert + engine inert),  plus zero dead-head payload,  as the burnout mass at Earth arrival.  This starts with a best guess for inert tank mass,  as well as for installed engine thrust level.  Apply the appropriate mass ratio to get Mars departure ignition mass.  The difference in ignition vs burnout mass is the propellant expended for the two burns of the return voyage.

The next step is the outbound voyage.  The Mars departure ignition mass,  plus the dead-head payload mass,  is the Mars arrival burnout mass.  Apply the appropriate mass ratio to get the Earth departure ignition mass.  The difference in ignition vs burnout mass is the propellant expended for the two burns of the outbound voyage to Mars.

The sum of the two propellant quantities is the total propellant for the round trip.  Divide this total propellant by ftank to find the total loaded tank mass.  The difference between loaded total tank mass and total propellant mass is the inert tank mass.  This resulting inert tank mass is what your guess for tank inert mass must converge to!  The best next guess is close to the last result.

Thrust divided by Earth weight is the vehicle acceleration gee estimate.  This is done at each of the 4 vehicle masses:  Earth departure ignition,  Mars arrival burnout,  Mars departure ignition,  and Earth arrival burnout.  Two of these are of real interest:  Earth departure ignition (min gees),  and Earth arrival burnout (max gees).  The other two conditions fall in-between. You must adjust your installed thrust level to achieve min gees.  Then iterate to convergence again on tank inert mass.

Max gees at Earth arrival burnout did not prove to be a problem,  but should fall under 5 gees for the most tolerable results.  Be sure you check for that outcome.

The last calculation sets up weight statements and estimated dV performance for the six propulsion types,  using the data already calculated.  The initial part of the weight statement is the vehicle buildup from payload and inert items to Earth departure ignition mass.  Subtracting the total outbound propellant gives the Mars arrival burnout mass.  Their ratio produces an outbound summed dV for both burns,  to be calculated for each type (for comparison to the initial summed requirement). 

That Mars arrival burnout mass,  less the dead-head payload,  is the Mars departure ignition mass.  Subtracting the return voyage propellant produces the Earth arrival burnout mass.  Their ratio produces a return summed dV for the two burns,  done for for each propulsion type (for comparison to that summed requirement).

The deviations of these weight statement dV’s from the required values reflect just how closely you converged your tank inert weights.  These should be only trivially off (by under 0.001 km/s = 1 m/s).  If you see bigger errors,  you didn’t converge your tank inert masses closely enough.  The effect of being “off” on min gee (as set by installed thrust level) is small,  when compared to the effect of being “off” on guessed tank inert mass.

At the very bottom of the weight statements are the vehicle payload fractions,  in both definitions.  One is the conventional definition:  dead-head payload mass / Earth departure ignition mass.  You probably should not consider anything under 0.2 for a practical colonization ship design.  Its inverse is Earth departure ignition mass / dead-head payload mass.  In that definition,  you probably should not consider anything over 5 for a practical colonization ship design. 

This limit (in either form) is inherently a very fuzzy judgement call.  But,  if dead-head payload mass is too small compared to Earth departure ignition mass,  the resulting design will be inherently very expensive to build and to operate,  just like with ocean-going transport when the cargo mass is small compared to the tonnage of the ship.

What I got for this study is given in Figures 1 and 2,  a two-part image of the completed spreadsheet worksheet page.  Of the six propulsion types,  four look reasonably-to-very attractive.  These are the derivative of NERVA,  some form of PBR,  and the two gas core concepts that do not require a huge waste heat radiator.  The as-tested NERVA falls short because its engine thrust/weight is too low and the resulting large engine inerts drive the vehicle inerts,  constrained by the large thrust level to achieve min acceleration gees.  The gas core with radiator falls short because of the gigantic,  heavy radiator.

Figure 1 – Image of Nuclear Thermal Spreadsheet Analysis,  Part 1

Figure 2 -- Image of Nuclear Thermal Spreadsheet Analysis,  Part 2

Near-term,  the higher Isp and engine thrust/weight of the derivative NERVA could be realized in a few short years,  to an engine prototype ready for flight test.  The PBR concept would take a few more years than that,  since no prototype engines were ever ground tested,  and some fundamental problems identified in testing of “Timberwind” components remain unresolved.  The gas core concepts would require several-to-many years to reach a flight-testable prototype,  since only very sparse lab-type feasibility demonstrations were ever done;  plus,  there is no guarantee of eventual success,  either.

My own recommendation would be to base an initial design around the derivative NERVA as lowest-risk option of acceptable benefit,  and plan on replacing it later with one of the non-radiator gas core designs,  should that development prove successful.

Figure 3 sketches a ship design concept based on the derivative of NERVA,  figured at 100 metric tons of dead-head payload delivered to Mars.  Volume of LH2 and a guess for tank L/D set the tank dimensions.  Everything else scales one way or another from that,  as a first approximation.  Everything about the weight statement and thrust level sizing is proportional to dead-head payload size.  Dimensions would scale as the cube root of mass,  provided that L/D ratios are preserved.

This vehicle rough-out delivers the same design dead-head payload to Mars as the proposed Spacex “Starship” design.  The differences are several:  this vehicle never lands on Mars (delivery to the surface is by unspecified other means),  this vehicle must make a full Mars arrival burn into low orbit (“Starship” only makes a final touchdown burn after an aerobraking direct entry),  and this vehicle returns all the way to low Earth orbit for reuse,  unrefueled.  It never needs to survive any sort of atmospheric entry. 

This design makes the round trip single-stage unrefueled.  The Spacex “Starship” is entirely one-way only,  unless and until it can be refueled on the surface of Mars from local resources.

There is enough payload section volume to support a crew of up to 15,  at about 300 cubic meters per person,  in addition to the volume occupied by the dead-head payload,  at a payload specific gravity averaging only 0.3.

Figure 3 – Sketch Layout of Derivative-NERVA Colonization Ship,  at the 100 Ton Payload Size

This result says a Mars colonization ship able to carry 100 metric tons of dead-head payload one-way to Mars,  and return to Earth with no payload,  all one-stage,  is not that large an item.  It is not large enough to spin for artificial gravity like a rifle bullet,  but it is large enough to spin end-over-end (like a baton) for artificial gravity.  At about 3.24 rpm,  there is about one full gee available in the payload section.  That spin rate is tolerable to untrained,  unacclimatized people,  for long-term exposure.

The insulation and meteor shielding is about a meter thick on the payload section,  meaning it can double as radiation protection.  If those layers of fabric average 0.20 effective bulk specific gravity,  that is some 20 g/sq.cm shielding mass,  adequate for solar flare events,  and offering some reduction of galactic cosmic radiation.  The insulation and tank shell thickness of the propellant tank section was assumed to be 0.1 m.  Engine section length was just a guess.

Key to this design as-sized is carriage of dead-head payload to Mars,  but not from Mars.  The return dead-head payload must be zero!  If not,  the propellant tank section must be significantly larger,  to the detriment of the payload fraction criteria.  Any crew and their life support must come out of that dead-head payload allowance (meaning near-zero crew on the return voyage).

These results look more favorable than the otherwise-comparable nuclear thermal option in the earlier study.  That is precisely because dead-head payload is only carried one-way in this study,  and it was carried both ways in the earlier study.  That is one huge effect.  But the trend from the earlier study applies here as well:  if we design for a farther destination than Mars,  the design won’t look so attractive in terms of the payload fraction criteria.

The restriction of zero dead-head payload on the return voyage is not as constraining as it first sounds,  when one considers the goal is building a colony with these payloads.  During that process there is little-or-nothing to ship home to Earth,  except information,  which is better sent electronically.  Later,  when an operating economy results in two-way trade,  one will need commerce ships,  not colonization ships.  But,  by the time that need arises,  significantly-better propulsion technology should have become available. 

Brief Result Summary:  The best near-term option of the six nuclear thermal approaches,  for a Mars colonization ship design,  is the derivative-NERVA nuclear thermal propulsion approach (Isp ~ 1000 s and engine T/W ~ 5).  For 100 metric tons dead head payload,  the initial ignition mass is about 500 metric tons.  That means for 1000 metric tons dead-head payload,  the sized ship will initially mass about 5000 metric tons.  For 2000 tons payload,  the ship will be around 10,000 tons,  etc.  This is restricted to orbit-to-orbit operation,  and to no dead-head payload on the return voyage.   Even the small 100-ton payload size is large enough to spin end-over-end for artificial gravity at near 1 gee and an easily-tolerated spin rate.  The payload section insulated design (if a meter of fabric layers) also inherently provides a fair amount of radiation protection.

Colonization Ship Study

I have gotten involved with some friends on the New Mars forums discussing what might be appropriate for very large colonization ships.  This kind of mission demands the delivery of very large payloads.  Doing this effectively requires a reusable ship.  That means you stage off (or jettison) nothing.

It is easy to run a rocket equation-based trade study that assumes a one-stage round trip,  that jettisons nothing.  Making it carry the same large payload on the return voyage simplifies the analysis,  but very likely over-penalizes the design.  But at this level of analysis,  that really doesn’t matter.

This is basically just a bounding analysis for screening candidate propulsion approaches to a Mars colony ship design.  I included nuclear explosion propulsion,  nuclear thermal propulsion,  ion propulsion,  LOX-LH2 cryogenic chemical propulsion,  and storable chemical propulsion.  

Update 9-13-19:  there is more than one kind of nuclear thermal rocket.  I took a closer look at 6 different nucear thermal rocket approaches,  and in a more nuanced way,  in "A Closer Look At Nuclear Thermal",  dated 9-13-19,  this site.

Spreadsheet Inputs

The spreadsheet inputs are highlighted yellow.  Payload delivered is common to all the designs,  and actually arbitrary,  but I thought 2000 metric tons might go a long way toward the beginning of a colony. 

Inert fractions vary with the propulsion selection.  I used data from Ref. 1 to set a realistic guess for the inert fraction,  of the nuclear explosion drive.  It is very high,  reflecting the massive pusher plate,  two-stage shock absorption system,  and the armored hull.

The Hall effect ion drive is based on existing Busek satellite thrusters already in service,  and modified to “burn” iodine,  something plentiful,  cheap,  and storable at low pressure.  Getting to an acceptable vehicle acceleration requires a very large thruster array and a nuclear power source in the multi-megawatt range.  I just guessed the inert mass fraction that might cover this. 

Because of the heavy reactor core and low engine thrust/weight achieved in the old NERVA nuclear thermal rocket development effort,  I used twice the typical chemical stage inert fraction as a “good guess” for the nuclear thermal inert mass fraction.  There is good data about this engine in Ref. 2.

Both the LOX-LH2 cryogenics chemical propulsion,  and the NTO-MMH storable-propellant chemical propulsion,  share the same “typical” stage inert mass fraction. 

Delta-vees for the Mars trip are for departing and arriving in low Earth orbit to/from a min-energy Hohmann transfer ellipse,  plus the corresponding delta-vees for arriving into and departing from low Mars orbit.  The same applies to the Ceres transfer,  except that the ship just matches Ceres orbital velocity about the sun instead of entering a “low orbit”.  This would be typical of many small main belt asteroids. 

For those types of propulsion in the order listed above (nuclear explosion,  nuclear thermal,  ion drive,  LOX-LH2 chemical,  and storable chemical),  my assumed inputs for Isp were 10,000 sec,  1000 sec,  3000 sec,  470 sec,  and 330 sec respectively.  Vehicle inert mass fractions were 0.50,  0.25,  0.10,  0.05,  and 0.05 respectively.   

All these dV’s were summed,  as required to do the entire mission single-stage.  The total orbital delta-vee (dV) to and from Mars is 3.84+1.83+1.83+3.84 = 11.34 km/s.  Impulsive-burn options need supply only that summed delta-vee with zero gravity and drag losses.  Long-burn ion must supply a lot more than that,  due to very large planetary and solar gravity losses.

All but the ion option were considered as "impulsive burn" and Hohmann min energy transfer,  with vehicle acceleration exceeding 0.1 gee to enforce that.  These used the unfactored sum of orbital dV's to and from Mars (orbit-to-orbit transport) as the mass ratio-effective dV for the rocket equation.  The spreadsheet input is factor equal to one. 

The ion option must spiral-out and spiral-in at the planetary orbits,  and accelerates to midpoint then decelerates to arrival on the transfer trajectory (a patched spiral about the sun).  Propulsion is sized for 0.001 gee to ensure that this kind of transfer is feasible.  To account for the planetary and solar gravity losses of the resulting months-of-burn,  I just doubled the orbital dV sum to 22.68 km/s.  For the spreadsheet,  this is factor equal to two.

For Ceres,  Earth departure and arrival dV is 5.24 km/s.  The orbit-matching dV at Ceres (arrival and departure) is just about 3.49 km/s.  That round trip sum is 17.46 km/s for all but the ion drive option,  unchanged by factor equal to one.  Using factor equal to two for ion drive,  that mass ratio-effective total is 34.92 km/s.

All 5 designs carried exactly the same 2000 metric tons of dead-head payload,  an arbitrary selection perhaps appropriate for a colony-type mission.  (I did not look at how to get that payload up to LEO,  or down from LMO,  that issue would be the same for all the candidates.)  This was done for Mars in a spreadsheet worksheet,  whose image is Figure 1.  All figures are at the end of this article.

Analysis Equations

Sum the round trip delta-vees,  and factor the sum for the mass ratio-effective delta-vee required of each propulsion type:  required dV = (factor)(sum of all 4 orbital delta vees),  where factor = 1 for impulsive propulsion (acceleration exceeding 0.10 gee),  and factor = 2 for long-burn ion propulsion (0.001 gee required).

Estimate the effective exhaust velocity from the specific impulse:  Vex, km/s = 9.8067 (Isp, s)/1000

Calculate the mass ratio required:  MR = exp(dV/Vex),  with both velocities in km/s

Calculate the propellant mass fraction:  Wp/Wig = 1 – 1/MR

Input an inert mass fraction Win/Wig (must be justified in some way as “realistic”)

Calculate the available payload fraction Wpay/Wig = 1 – Win/Wig – Wp/Wig  (must be positive to be even theoretically feasible)

Input the delivered dead-head payload Wpay,  metric tons (arbitrary,  but should be realistic)

Calculate the ignition mass Wig,  metric tons:  Wig = Wpay/(Wpay/Wg)

Calculate the inert mass Win,  metric tons:  Win = Wig*(Win/Wig)

Calculate the propellant mass Wp,  metric tons:  Wp = Wig*(Wp/Wig)

Calculate the ignition to payload mass ratio:  Wig/Wpay = (Wig, m.ton)/(Wpay, m.ton)

Results Obtained

Results for Mars:  nuclear explosion drive 5118 metric tons at ignition with ignition/payload 2.56:1 (see Figure 2).  Nuclear thermal 30,945 metric tons at ignition with 15.47:1 ignition/payload (see Figure 3).  Hall effect ion drive 5516 metric tons at ignition with ignition/payload 2.76 (see Figure 4).  LOX-LH2 56,486 metric tons at ignition with ignition/payload 28.24 (see Figure 5).  Storable chemical utterly infeasible with a negative payload fraction available (see Figure 6).

The nuclear explosion drive offers the lowest ignition/payload ratio going to Mars at 2.56:1,  based on the old 1950's shaped-charge fission device technology.  This would be a very tough ship design,  probably usable for a century or more,  and likely tough enough to aerobrake,  reducing the load of bombs in favor of more payload.  Its stout hull and huge pusher plate are effective radiation shields.

The ion propulsion offers the next best ignition/payload ratio going to Mars at a very comparable 2.76:1,  which to be practical would require its thrusters operating on something cheap,  plentiful,  and storable-as-a-condensed-phase (at very low pressure),  like iodine.  This would be a relatively gossamer structure unable to survive aerobraking,  and it would likely also have a limited service life.  Radiation protection would have to be added.

Two of the others (nuclear thermal and LOX-LH2),  while theoretically feasible,  are nowhere close in ignition/payload ratio going to Mars.  These are unaffordable “Battlestar Galacticas” for any reasonable payload delivery aimed at colonization.  And the storable chemicals are just infeasible in any sense of the word for a Mars colonization ship,  simply because there is a negative payload fraction available,  once propellant fraction has been determined,  and with a suitable inert fraction input.  It simply cannot do the mission single stage.

I think you can look at the ignition/payload mass ratio to judge whether-or-not a given propulsion system might serve as a practical way to build a colony ship.  This value needs to be no more than about 5 or thereabouts,  in order not to build an unaffordable “Battlestar Galactica”.  This is a “fuzzy” boundary,  dependent upon how much you think you can afford.

The same sort of analysis applies to other destinations.  You just need an appropriate list of orbit-to-orbit delta-vees,  and the same list of realistic guesses for inert fractions.

Results for Ceres:  I added a worksheet to the same spreadsheet for a colony-type ship to Ceres,  as “typical” of the asteroid belt.  Those spreadsheet results are shown in Figure 7.  Figures 2 – 6 also show the Ceres results (as well as the Mars results). 

The only feasible choices for Ceres colony ships were nuclear explosion propulsion and nuclear-powered electric propulsion.  It’s the same basic calculation,  just with somewhat bigger delta-vees.  The nuclear thermal and both chemical options simply had fundamentally-infeasible negative payload fractions available.  They simply cannot perform the mission single-stage.

The same general outcome choices obtain for Ceres as for Mars:  your nuclear explosion drive ship is quite robust,  promising a long service life,  while the ion ship is rather flimsy.  For this main belt asteroid application,  the ignition to payload ratio is also substantially more favorable for the nuclear explosion ship  (2.97),  vs the ion ship (4.87).

Conclusions

The trend here is clear:  the further out you go with a single-stage,  round-trip colony ship,  the more the ignition/payload ratio is going to favor nuclear explosion propulsion as the more affordable option.  Radiation protection needs will also favor the shielding effect of the stout hull required of the nuclear explosion drive.  Bigger also favors ease of incorporating spin “gravity”. 

References

#1. George Dyson,  “Project Orion – The True Story of the Atomic Spaceship”,  Henry Holt,  2002.

#2. David Buden,  “Nuclear Thermal Propulsion Systems”,  Polaris Books,  2011.

Figure 1 – Spreadsheet Image:  Mars Colonization Ship

Figure 2 – Results Summary for Nuclear Explosion Propulsion

Figure 3 – Results Summary for Nuclear Thermal Propulsion

Figure 4 – Results Summary for Iodine-Fueled Hall Effect Ion Propulsion

Figure 5 – Results Summary for LOX-LH2 Chemical Propulsion

Figure 6 – Results Summary for NTO-MMH Chemical Propulsion

Figure 7 – Spreadsheet Image:  Ceres Colonization Ship

Bittersweet Event

Final Update 11-17-20:  Yesterday I closed the deal to sell this plane to a new owner.  He will take excellent care of it,  and put it to a useful purpose training student pilots.  That was very important to me,  as this plane has enormous sentimental value.  

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In a posting dated 18 January 2014 and titled “Super Red-Letter Event”,  I described inheriting my Dad’s plane and learning to fly in it.  In the last couple of years,  I have had some serious health problems that prevented me flying,  culminating in a mild stroke that has effectively ended my flying. 

This aircraft needs to be flown,  it will deteriorate if just left in storage.  It really needs an owner who can take full care of it,  which inherently includes flying it,  something I can no longer do. 

This aircraft is truly a vintage craft;  I typically referred to myself flying it as the “antique flying the antique”,  but it is not worn out!  It is a strikingly-classic design,  and draws very positive comments from observers quite often.  That rounded vertical tail is nothing at all like anything you can buy today.  That really attracts attention,  not to mention it being a tailwheel design.

It is fun to fly,  and easy to fly,  although you must be tailwheel-qualified to fly it (something rare today).  Tailwheel instructors are also now rare,  but they all claim a tailwheel-qualified pilot is actually a better pilot,  precisely because he/she must actively “fly” the plane,  even on the ground,  from the time the prop starts turning,  until it stops turning back at the hangar.  I agree with that opinion.  It’s not hard,  but it does demand very close attention to everything,  even on the ground.  Especially the wind.

This isn’t modern high-speed flying glued to a “glass cockpit” that does almost everything for you.  It is flying as it was done well over half a century ago:  looking out the window and using a real chart and a VOR to navigate,  all the while monitoring round dial-type gauges on the instrument panel.  It’s not hard,  but you need practice at it,  to do it effectively in the air.

Your piloting skills actually stay sharper if you fly that way.  This is real “stick-and-rudder” stuff!  Yet this particular plane flies fast enough to be attractive for travel:  it cruises near 110-115 mph.  Day VFR / “stick-and-rudder” is easy in it,  and the most fun of all.

Below is my writeup (complete with a photo) describing this airplane for prospective buyers.  If this kind of flying appeals to you,  and you might want to buy this plane,  do please contact me.  I put the contact data in the writeup.

GW

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writeup  (updated slightly 9-1-19)
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4-Seat “Old-Time” Cessna for sale.  Owner can no longer fly due to health issues.  This aircraft needs an owner that can fly it,  so that it does not deteriorate just sitting in a hangar.

1952 Cessna 170B,  S/N 25336,  registered as N2794D,  Continental C-145 engine S/N 8188-D-3-2,  McCauley 1A170 propeller,  S/N 71827.  Original owner’s manual. 

Tailwheel aircraft,  requires tailwheel endorsement!  Standard category,  not light sport!

Aircraft has been hangered all its life since being rebuilt from salvage with a new engine at an estimated 2200 airframe hours,  and returned to flight in 1983 (logs date from then).  All metal wing,  3-position flaps,  dial-type gages on instrument panel.  Cruises at about 115 mph.  Easy to fly.  Insured for $41,000 with Avemco (their professional minimum valuation).

Total time in service 2716.7 hours as of last annual.  Engine time in service since last major overhaul 535.4 hours as of last annual.  This is a relatively low-time airframe,  and a fairly low-time engine!  Compliant with all AD’s as of last annual.  Date of that last annual:  January 2016.

Aircraft was flown less than an hour,  in a short ferry flight,  since that last annual.  Stored in T-hangar at McGregor airport ever since,  and also hangered ever since current owner took possession.  Was also always hangered before that,  dating back to its rebuild and return-to-flight.  

Rudder trim tab needs adjustment after last repair:  replace broken tailwheel leaf spring and repair associated sheet metal damage to rudder and elevators.  This was done during calendar year 2015,  preceding that last annual. Ferry flight revealed a need to add some “R turn” at trim tab on rudder.

Aircraft has proper radio,  VOR,  and transponder,  but lacks ADS-B “out”,  which is soon-to-be-required for operation near towered airports from 2020-onward (there are now less-expensive solutions for this).  Will need a current annual to return to flight.  

Excellent appearance:  paint good.  Upholstery good,  except pilot and copilot seat covers are worn,  but still very serviceable.  Carpet padding and seat foam padding needs replacing.

Contact owner at:

Gary W. Johnson,  PE,  PhD

5886 New Windsor Pkwy

McGregor,  TX  76657

254-840-9629

gwj5885@gmail.com

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end writeup
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Update 9-1-19: 

Looking at some recent Trade-A-Plane ads,  asking prices depend on the model year.  The plot following the photo is that Trade-A-Plane data.  That trend says a 1952 model is worth just about $50,000,  “all else being equal”.  The pluses and minuses are a “wash”,  so all else really is equal,  or better.  Therefore $50,000 is my asking price.  The minimum I would consider is the insurance evaluation $41,000.  

UT Austin Gang

This photo was taken at the last gathering of this group earlier this month.  These are my closest friends from my UT student days,  and their spouses. 

Front row left-to-right are Laurie Mahaffey,  my wife Ellen,  myself,  Janie Caldwell,  and Mike Caldwell.  Back row left-to-right are Mike Mahaffey,  Jesse Boultinghouse,  Dinah Boultinghouse,  Sissy Moore,  Jack Moore,  and (one behind the other) Susan and Mike Brands.

I met Mike and Laurie Mahaffey as a UT freshman.  They were together even then.  Mike and I were in aerospace engineering,  Mike in the avionics specialty,  and me in aerothermodynamics and propulsion.  Mike has yet to retire from a long career flight testing equipment.  I left aerospace upon plant closure,  after a long career doing mostly rocket and ramjet development work.  After that,  I did mostly teaching.  It was Mike who set his phone camera to take this picture all by itself.  

Jesse Boultinghouse was in chemistry.  He and Dinah got together long after we all finished school.  Jesse works for the state in water quality.  Jack Moore was a physics major.  He and Sissy met long after we finished school.  Jack had a career in nuclear weapons work,  among other things.

Mike Brands was in mechanical engineering and naval ROTC.  He knew Susan back then,  but it was only in recent years that they got together.  Mike had a long career in the navy,  then some civilian things.  Susan was the sister of another good friend Terry Forman,  also a physics major,  who lives on the east coast. 

Mike Caldwell was a business major at UT,  and was in army ROTC.  He and Janie got together about the time I finished graduate school,  I think it was.  Mike had a long career in logistics in the army,  and consulted for the army in it,  for some time after he retired from the army.

All these men were dormitory and later apartment mates with me,  during my undergraduate and graduate years at UT,  except Mike Mahaffey.  Mike was in a lot of the same classes with me,  especially in undergraduate school. 

Two other good friends from UT are not pictured here,  because they are deceased.  They are Roger Prior and Terry Boone,  both dorm mates.  May they rest in peace. 

Anyhow,  this group as pictured gets together every several months,  for good food and chat.  

Trump 2020? Nope!!!

This image actually speaks for itself,  although one of the nuances in it may be lost on most folks.  Mr. Trump did not notice this altered image of the presidential seal behind him.  The person who put this there,  got fired for doing it.  I got this off the PBS NewsHour website,  not “social media”.  It is credible.

Everybody notices the golf clubs,  and the connection to what Mr. Trump loves to do at Mar El Lago.  In the real presidential seal,  these are arrows,  contrasting with the olive branches in the other talon.

What most folks may miss is the double-headed eagle.  The only nation that uses the double-headed eagle symbol is Russia.  So this altered seal really took two swipes at the President,  not just one. 

Rather clever,  really.

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Update 8-3-19:  it has come to my attention that there is a second level of the swipe against Trump about the Russian connection.  This is very hard to see,  especially in my copied photo,  even if you enlarge it.  But,  the motto in the banner has been altered from "E Pluribus Unum" to "45 Es Un Titere",  which translates to "45 is a puppet".

The revised motto is Spanish,  not Latin,  which connects to the border crisis manufactured by bad policies and neglect.  And,  in yet another level of swipe,  the olive branch in the other talon has been replaced by a wad of cash,  although this,  too,  is difficult to see in my copy of the photo. 

The news stories indicate that the artist who drew this is a former more-or-less independent voter who has become very disenchanted with Trump.  He drew it for emotional catharsis,  and did not intend it to be used this way. 

Somebody else searching for a presidential seal to project during a Republican event found this on the internet,  more probably by mistake than malicious intent,  and got fired for using it. 

The evils of political extremism show in this incident in two ways:  (1) the artist has gotten hate mail for creating this,  and (2) for the most part,  the individual who used this at the event has been automatically presumed to an intentional political saboteur. 

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

Why would I post this?  Well,  let’s just say I have good reasons not to be a Trump supporter,  which makes this image even funnier than otherwise.  Read on to understand why,  being warned this part is not funny at all.

Untruthfulness

Mr. Trump is wrong in his claims very nearly 100% of the times that he has been fact-checked.  That track record suggests you should believe nothing he says. How is having such an egregious chronic liar in the White House a good thing for America?  

The Economy

Mr. Trump has sort-of-maintained the economic recovery that began under Mr. Obama,  in spite of all the damage he has been doing with his trade wars and tariffs.  You don’t have to start a trade war to renegotiate a trade agreement.  This cannot last:  the economy will crash,  and it will happen on his watch,  directly attributable to what he has done with those trade war things.  How are such idiotic policies good for America?

Bad Immigration Policies

Mr. Trump has instituted immigration policies that by his own admission,  and the admission of others in his administration,  intend to dissuade other potential immigrants from ever coming,  by grossly mistreating those already at our border.  This continues,  in defiance of a court order to cease and desist.  How is that not a form of both racism and state-committed terrorism?  How is such a policy of racism and terrorism good for America? 

Former president Andrew Johnson was impeached for racism in 1865,  although he was not convicted in the senate.  Racism is therefore an impeachable offense. 

We have fought against state-committed terrorism since World War 2 (the Gestapo and the SS as examples).  This evil on our border continues today under the Trump administration because not yet enough good people are standing up against it.  If we fought wars against state-committed terrorism,  then how is its commission by our own government not an impeachable offense?

Lies About the Mueller Report

There are still a lot of lies being told about the Mueller report,  most (but not all) coming from the GOP.  Unfortunately for the country,  they seem to have prioritized political advantage and getting re-elected over the good of the country.  If you read the report for yourself,  you can see through the lies,  and see what it really says,  despite the redactions in the publicly-released form.  I did.

The political lies about Mueller’s report continue,  hoping no one will read for themselves the truth of the matter.  This is party politics at its worst. I recommend you hold them accountable by not re-electing them,  next time.

Volume 1 documents conclusively (1) that the Russians interfered in our 2016 election (and how they did this),  (2) an eager willingness of Trump and his campaign to cooperate with the Russians who were wanting to help get him elected,  and (3) an unwillingness of Trump and his campaign to report these illegal attempts by the Russians to interfere (note that foreign participation in a US election effort is a federal crime). 

It does clear Trump and his campaign of conspiring with the Russians (ahead of time) for them to commit this interference crime to benefit him. THAT is the only “exoneration” anywhere in Mueller’s report.

Volume 2 documents some 10 instances of obstruction-of-justice on the part of Trump and his minions,  any one of which is likely an impeachable crime.  Mueller chose not to indict,  or to definitively-conclude that crimes were committed,  based on a Department of Justice memo claiming that sitting presidents cannot be indicted. 

But he documented the crimes,  and he intended Congress to follow up on them. Which they are,  at least in the Democrat-controlled house.  He specifically said that while he did not officially accuse Mr. Trump of obstruction of justice,  he did NOT exonerate Mr. Trump of it,  either.

Siding With Putin Instead Of His Own Intelligence Agencies or Allies

The intelligence agencies and the Mueller report document that the Russians interfered in the 2016 election,  and just how they did this crime.  Mr. Trump has repeatedly in public taken as truth Mr. Putin’s denial of this. 

In addition,  Mr. Trump cozies-up to Putin,  Kim Jong Un,  and China’s leader,  while insulting our allies or chastising them over money spent on NATO.  The Russians,  under two different governments (Soviet and Putin) have tried without success for about 7 decades to weaken NATO.  Now our NATO allies have doubts we will come to their aid if needed,  and it is Mr. Trump who did this weakening of the alliance.

How is Russia under Putin not a hostile power?  How is either problem not “aid and comfort to the enemy”?  Read the definition of treason in the Constitution for yourself.  How are these two things not at least bordering upon treason of the aid and comfort type?  How is not holding Mr. Trump accountable good in any way for America?

What Shall We Do?

It’s getting close to the 2020 election now.  It may be too close for a real impeachment proceeding,  especially since the GOP-controlled senate still appears entirely unwilling to convict.  THAT dereliction-of-duty on the part of the GOP senators is another topic,  my point here is that Mr. Trump does NOT need to be president anymore.  Whatever good he might possibly do (or have already done) is far outweighed by the evils I have listed.

To that end,  the short form sound bite is:

“Dump Trump 2020”

The longer-form rationale is:

“Doesn’t matter who the Democrats run,  how could you possibly be worse off?”

If the house chooses to impeach between now and the election,  it should be aware that a non-conviction in the senate will motivate Trump supporters to turn out and vote.  This is not because of facts,  but because of their belief system (that Trump is good for America despite his faults).  True believers rarely respond to facts. Only something truly egregious could overwhelm that belief.

A better strategy might be to use the extra power of “impeachment-related proceedings” to uncover the tax records of Mr. Trump. 

These will probably reveal that the bulk of his investor capital since the casino bankruptcy is coming from the same Russian banks that symbiotically keep Mr. Putin in power in Russia.  If Putin really wants something of Trump,  the banks will demand it of him.   He is “in hoc” to them.

I do not know this to be a fact,  but it seems very likely to me,  because of how desperate he seems NOT to have those records made public.  It’s a “guilty dog barks loudest” sort of thing.  Not proof,  but good cause to investigate.

THAT is the type of egregious-connection-to-Russia that might open some Trump supporter eyes,  as to who and what he really is (a clear-and-present danger to our democracy).  A revelation like that would likely cost him the election,  if not a senate conviction. 

Just Mooning Around

Updates:

Addendum A added below 7-20-19
Addendum B added below 8-19-19
Addendum C added below 8-19-19

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Tuesday July 16 is the 50^th anniversary of the launch of Apollo 11 to the moon.  Saturday July 20 is the actual 50^th anniversary of the landing.  Those of us old enough,  remember where we were and what we were doing,  when the news of the landing broke.

I was attending the US Naval Academy at the time of that landing.  The moonwalk took place in the wee hours of the morning,  in that time zone.  We otherwise-incommunicado midshipmen were awakened and taken to see the televised event of history being made. 

Space Race

This truly historic event was the culmination of a cold war space race with the Russians,  that started with their orbiting of Sputnik in 1957.  I remember watching that first satellite in the evening sky,  and I have watched a great many since.

The US seemed to play catch-up for several years afterwards,  as the Russians orbited the first man in space,  took the first spacewalk,  flew the first woman in space,  and flew the first multi-person crews.  Until 1958,  the US did not have a NASA at all.

With some exceptions,  our effort was a cost-is-no-object,  take-some-serious-risks “crash program”.  The Christmas 1968 Apollo 8 mission around the moon was an example of taking some very serious risks.  It was really the first manned test flight of the Saturn-V rocket.

But it worked.  Their giant rocket had failed too many times while ours did not,  which is what led to Apollo 11 and the other 5 landings.  The goal was to beat the Russians to the moon,  and it was accomplished.

Consequently,  what we did was less about science and exploration,  and more about experimental test flight work.  This was true from the very first Mercury shot (Alan Shephard in 1961) onward.  All but one of our moon-walking astronauts was an engineering test pilot. 

Only Harrison Schmidt on the very last landing was originally educated as a geologist.  Our astronauts were trained to set up experiments,  and to collect certain types of rocks,  so that other scientists back home could do the actual science work later.  That work is still going on.

Why We Never Went Back All These Decades 

There was no longer any “race” to win.  Because there was no “race” to win,  there was no support for spending large amounts of government money.  Typically,  the funding for science and exploration projects here on Earth was far smaller,  and most of it was historically non-governmental.

Government funding is very highly politicized.  John F. Kennedy committed us to a moon landing before 1970,  but he was not a fan of the space program,  until he witnessed a Saturn-V engine test.  It was Lyndon Johnson who was the real believer in a space program,  and he convinced Kennedy that beating the Russians to the moon was really something possible and worthwhile.  

The presidential leadership continuity that got this 8-year effort accomplished was mainly because Johnson succeeded Kennedy.  Richard Nixon,  who was not a big supporter of the space program,  succeeded Johnson,  and he killed the Apollo moon program at the 6^th landing (Apollo 17). 

We were supposed to have flown another 5 landing missions all the way through Apollo 22.   Leftover unused rockets and capsules are why there are so many of these items on display around the country today. 

The big-ticket manned space items since,  have been the Space Shuttle and the International Space Station.  These have been marvelous things,  to be sure,  but they are as much about “workfare” for the businesses that do this,  and porkbarrel politics for the congressional districts where they are located,  as anything else.

Why We Should Go Back (And Farther)

In the long run,  it’s not about winning some race,  and it’s not so very much about doing pure science just for the sake of knowledge.  It’s about exploration of the unknown,  something hard-wired into humans.  In centuries past,  this was exploration of the unknown parts of the Earth. 

“Exploration” is a really an emotionally-loaded code word.  What it truly means is you go there to find out what all is there (resources),  and where exactly it is (how hard to obtain).  Then you stay a while to figure out how to use what you found,  to cope with living in the local environment.

Unless you do that correctly,  there is no possibility of future settlements and the associated future economies!  There is no way to accomplish anything else,  except just the act of going there and returning (which is the bulk of what Apollo accomplished).

Is there anything worthwhile to accomplish out there?  Yes,  definitely! 

In the longer term,  there are those future off-world settlements and the associated future economies.  I cannot tell you the details of how this might benefit us,  because it has yet to be done.  But it has always proven beneficial in prior centuries here on Earth.

In the shorter term,  there are the possibilities of space resource businesses,  and of planetary protection against rogue asteroid and comet impacts.  There is simply no better reason for continuing both unmanned and manned space programs than finding ways to protect the folks back home!

Change in Approach is Required

This cannot be done properly as a “NASA-does-it-all” set of projects.  It is likely not possible for the US to do it all alone,  either.  This cannot be done as a “crash program”,  because that is far too expensive.  But exploration is risky,  you cannot be too risk-averse,  either.  

Ethically,  you do have to address the known risks as best you can,  because we have learned to our chagrin that there is nothing as expensive as a dead crew!  If addressing those risks means you design vehicles and hardware different from anything ever seen before,  then so be it.  See addendum A below.

This definitely requires thinking way outside the boxes where we have been before,  with Apollo and Space Shuttle and ISS.  This is going to require a far stronger and more equal partnership between the government and the companies in the business.  And it requires looking far beyond the traditional contractor base.  Pork-barrel politics-as-usual and big-corporate “workfare” simply have to go away.  See addendum B below.

My Suggestions for the Near Term

Establish a continuous human presence on the moon,  the first item.  Start small and expand it slowly over time. Do the “exploration” thing right,  this time.

Send humans to Mars as the fulfillment of a dream centuries old,  probably the second item.  When we go,  do the “exploration” thing right,  from the very first landing. 

But,  any vehicle capable of taking crews to Mars can take a crew to near-Earth asteroids and comets.  Visit those asteroids and comets and properly explore them,  in order to learn how to defend against their impacting Earth. 

That’s the third item,  but it is just as easily done,  and at least as important,  as going to Mars.  Maybe we do them at the same time.  See addendum C below.

Update 7-20-19:  An edited version of this article appeared as a guest column today in the Waco "Tribune Herald".

Addendum A 7-20-19:  Ethically and Responsibly Addressing Known Risks For Spaceflight

In the main article,  I said we are ethically bound to address known risks of spaceflight as best we can,  but did not explore that topic.  Here are the details of that topic.

There is a whole long list of safety risks associated with any sort of spaceflight.  Three come to mind as the most truly credible risks:  (1) reliability of,  and escape from,  spacecraft and booster rockets,  (2) microgravity diseases,  and (3) exposure to radiation. 

The first one has cost us three American crews totaling 17 people (Apollo 1,  shuttle Challenger,  and shuttle Columbia).  Each caused a year-or-more stand-down,  and very expensive investigations,  plus very expensive changes.  The two shuttle losses were ultimately caused by bad management decisions.  This is why I claim “there is nothing as expensive as a dead crew,  especially one dead from a bad management decision”. 

Making spaceflight more safe from a reliability standpoint is also something we already know how to address!  This takes careful design allowing for failure modes,  redundant systems,  and copious verification testing.  Mitigation efforts will never be perfect,  but they can be quite good.

The other two have been long studied in low Earth orbit,  where microgravity exposure is inherent in everything we have done there,  and radiation exposure is somewhat more than on Earth’s surface,  but less than outside the Van Allen radiation belts,  and far less than inside the belts themselves. 

               Microgravity Diseases

Microgravity has proven to affect the human body in a variety of expected,  and unexpected,  ways.  The longer one is exposed,  the worse the various diseases become.  Beyond the bone decalcification and muscle-weakening we expected,  there are also degradations of the heart and circulatory system,  degradation of vision from eye geometry changes due to the fluid pressure redistribution,  and immune system degradations.  No doubt more will be discovered,  as that has been the trend. 

The longer exposed,  the longer it takes to recover upon returning home,  with full recovery actually still in doubt for some of the effects. The practical time limit seems to be only a bit more than a year.  Usual practices on ISS call for 6 months to a year’s exposure at most.

For operations anywhere in Earth-moon space,  one-way flight times are in the “4 days to a week or so” class.   Out of 6 months-to-a-year allowable,  that leaves at least several months available “on site” anywhere in Earth-moon space to do whatever needs to be done,  before too-serious ill effects occur. 

We do not know if low gravity,  such as the 16% on the moon,  is therapeutic enough to reverse the effects,  or at least extend the feasible exposure time.  That is because we never built the spinning space stations in low Earth orbit,  in which to experiment with different levels of artificial spin gravity.  In hindsight,  that lack was rather stupid of us,  now wasn’t it?

We do know that something near one full Earth gravity (one “gee”) is therapeutic,  because that is what we evolved in.  So,  until we know better,  any artificial spin gravity schemes need to supply very near one gee,  in order to obtain the full Earthly benefits that we know work. 

Destinations outside of Earth-moon space are very much further away:  one-way travel times range from near 6 months to multiple years.  This is pretty much outside the limits of microgravity exposure that we have established on ISS (no more than a year,  only 6 months preferred). 

Mars is 6-to-9 months away one-way,  and we do not know how therapeutic its lower gravity (38%) really is.  Other destinations are further away still,  and all those we can reach are even lower gravity than Mars.  That situation says quite clearly that we need to provide artificial gravity (no matter how inconvenient that might otherwise be) at something near one gee (until we actually know better),  during these one-way transits,  in order to best preserve the health of the crews. 

Ethically,  you simply cannot argue with that conclusion,  no matter how inconvenient for design purposes,  or for total mission cost purposes.

               Supplying Artificial Gravity

There is as yet no such thing as “Star Trek”-type artificial gravity.  The only physics we have to serve that purpose is centrifugal force.  You must spin the vehicle,  to generate centrifugal force as an equivalent to gravity.  If the spin rate is low,  then Coriolis forces (something everyone experienced on a merry-go-round) become less important,  and so fewer folks can tell the difference between this and real gravity.

The physics says that the acceleration you feel is proportional to the radius of spin and to the square of the spin rate.  The actual physics equation says

               a = R w²  where a is the acceleration,  R the spin radius,  and w the spin rate

Another form expressed in gees,  and not absolute acceleration units is

               gees = 1.00 * (R, m / 55.89 m) (N, rpm / 4 rpm)²

Experience with spin rates says that normal untrained and unacclimatized people can tolerate 3 to 4 rpm immediately,  for long-term exposures,  without getting motion sick.  People extensively trained might tolerate higher spin rates in the 8-12 rpm class without getting motion sick from long exposures.  Still-higher spin rates (16+ rpm) induce blood pressure gradients head-to-toe that are just unacceptable for long term exposures. 

The upshot of that is that the required spin radius (half of a crucial dimension of the craft you must build) will be about 99-56 meters at 3-4 rpm,  and about 14-6.2 meters at 8-12 rpm.  These are inconveniently large dimensions,  at least at first glance.  

3-dimensional objects typically have 3 axes.  About these axes these objects have a property called “moment of inertia” that relates to spin dynamics.  Usually,  higher moment of inertia correlates with larger dimension along a perpendicular axis to the spin axis. 

There are two (and only two) stable spin modes for most objects:  about the axis for highest moment of inertia (longest dimension),  and about the axis for lowest moment of inertia (shortest dimension).  The first case is exemplified by a baton twirler’s spinning baton,  and the second case is exemplified by a spinning bullet or artillery shell.  There are no other stable modes of spin.  See Figure A-1.

Figure A-1 – Modes of Spin

Clearly,  building a “spinning rifle bullet” 112 m in diameter at 4 rpm for one full gee at its outer girth is not so very feasible.  But spinning a smaller-diameter “something” that is 112 m long,  end-over-end at 4 rpm,  for 1 gee at each end,  would indeed be a feasible thing to attempt. 

We already know a lot about the transient dynamics of spinning rigid objects,  something important for spin-up and spin-down,  as well as for applying thrust while spinning.  This got started balancing steam locomotive wheels more than 200 years ago.  It continues today balancing wheels and tires on cars and trucks.  There would be no engineering development work to design a long,  narrow spacecraft that spins end-over-end for artificial gravity.  There would only be proving-out the specific design before we use it.

The most-often-proposed alternative is a cable-connected structure,  because it is conceptually easy to reel-out long cables between two small objects.  Cables only support tension loads,  not compression,  nor side-directed loads.  Transient dynamics for spin-up and spin-down,  and especially for applying thrust while spinning,  are incredibly complex and still not very well-known.  This is not something we have much experience with,  at all.  So there is a huge engineering development effort needed,  beyond just proving-out the actual design to be applied.

What this really says is that the preferred near-term spacecraft design is a long and rigid,  more-or-less cylindrical shape,  to be spun end-over-end,  baton-style.  This will generate varying gee from a maximum near the ends,  to zero at the spin center.  That is very likely the lowest-weight rigid-body design for any given spin condition,  and it is consistent with our long history rigid-body spin experience,  thus eliminating the huge engineering development effort required of cable-connected systems. 

We know that microgravity vs gravity has no impact while prone sleeping,  or else Earthly bed rest studies would not be a decent surrogate for some in-space microgravity effects.  That means you can put the sleeping quarters in the low gravity section of the spacecraft near the spin center,  and just put the daily workstations in the full-gravity sections of the spacecraft near the ends.  See Figure A-2. 

Figure A-2 – Why Selecting Baton Spin Mode Is Wisest Choice

               Radiation Hazards

There are basically three radiation hazards to worry about:  galactic cosmic rays (GCR),  solar flare events (SFE),  and the Van Allen radiation belts about the Earth.  All three are atomic or subatomic particles,  just at different speeds and quantities.  The threats they pose are location-dependent.

GCR is a very slow drizzle of really high-speed particles moving at a large fraction of the speed of light.  Particles that energetic are very difficult to shield against,  because they penetrate deeply into shielding material,  and quite often create “secondary showers” of other harmful radiation when they strike the atoms in the shield material.  If the shielding atoms are low atomic weight,  the secondary shower effect is greatly reduced.

GCR comes from outside the solar system.  Its quantity is affected by the solar wind,  in turn affected by the sun’s sunspot cycle,  which is about 11 years long.  The solar wind is stronger when sunspots are active,  making GCR lower in the vicinity of the Earth-moon system at that time.

From NASA’s radiation effects website,  I obtained these values that apply in the vicinity of the Earth-moon system.  GCR maximizes at about 60 REM per year when the sun is quiet,  and minimizes at about 24 REM per year,   when sunspots are most active.  To “calibrate” what may be unfamiliar units,  the natural Earthly background radiation is about 0.3 REM per year,  and a lethal dose would be 300 to 500 REM accumulated in a “short time” of hours to weeks.

The NASA astronaut exposure standards are set at about twice the levels allowed for Earthly nuclear workers.  Those NASA standards are no more than 50 REM per year,  no more than 25 REM in any one month,  and a career limit that varies with age and gender,  but maxes-out at no more than 400 REM accumulated over a lifetime.  These are predicated upon a single-handful percentage increase in the likelihood of late-in-life cancer.  

Clearly,  with a very modest shielding effect,  GCR is not the “killer” it is often portrayed to be.

SFE (solar flare bursts) are different.  They are much lower speed particles,  much easier to shield,  but there is an incredibly-huge flood of them.  They come in very-directional bursts from the sun at erratic intervals.  There are usually more of them during times of active sunspots,  but they can indeed happen when the sun is quiet.  They come at irregular intervals measured in several months.

The intensity of a burst can vary from tens of REM over a few hours,  to tens of thousands of REM over a few hours.  Obviously,  for unshielded persons,  the great bulk of events like this would be fatal doses,  and it is an ugly death.  There was a fatal-level event in 1972 between the last two Apollo missions to the moon,  and a low-intensity (non-fatal) event during one Apollo mission to the moon.

We had chosen to ignore this SFE threat during Apollo because the short duration of the missions (at most 2 weeks) was small compared to the typical interval between events.  Had a large one hit an Apollo mission,  the crew would have died in space.  As it turns out,  this was not a good assumption.
 

For an extended or permanent return to the moon,  shielding is obviously imperative.  On Earth,  we are protected from these SFE’s (and the GCR) by both the Earth’s magnetic field and its atmosphere. 

These are a very real threat anywhere outside the Earth’s magnetic field.  In low Earth orbit,  we are protected only by the magnetic field,  and the background exposure is higher than on Earth,  but much less than beyond the magnetic field.

The Van Allen belts are regions of these radiation particles trapped in the Earth’s magnetic field.  The intensity is lethal on a scale of days-to-weeks,  but tolerable on a scale of hours-to-a-day-or-so.  The inner boundary is not sharp,  but generally considered to become a problem at about 900 miles altitude,  and extending many thousands of miles out from the Earth. 

The exception is the “South Atlantic Anomaly”,  where the inner side of the Van Allen belt dips down locally to low Earth orbit altitude (100-300 miles).  Satellites and spacecraft in high-inclination orbits inherently pass through the South Atlantic Anomaly every several orbits.  The ISS does encounter this threat,  it being short “flashes” of exposure that accumulated over time still fall within the astronaut exposure standards.

Spacecraft traveling to the moon or elsewhere must transit the Van Allen belts.  Because of the potential for lethal exposure if you linger within them,  such transits must be made quickly.  Apollo did this,  transiting within only several hours.  Given the state of today’s electric propulsion technology,  this rules out using electric propulsion for people to leave Earth orbit for the moon or elsewhere,  because the spiral-out time is measured in months.  That would be lethal exposure,  even with some shielding.

               Passive Shielding

The same NASA radiation site has data regarding the shielding effects of typically-considered materials.  Those are hydrogen,  water,  and aluminum.  Mass of shielding above a unit exposed area turns out to be the correlating variable,  and 15-20 g/cm² seems to be enough to generally address the worst SFE. 

Hydrogen has the lowest density,  requiring the thickest layering,  but also has the least secondary shower potential against GCR. 

15-20 cm of water is 15-20 gm/cm²,  same shielding effect as a really thick layer of hydrogen.  Water molecules are still light enough not to have much secondary shower risk. 

Aluminum would be the thinnest layer,  but with the greater secondary shower effect.  However,  of the practical metals,  its atoms are the lightest,  and this secondary shower effect is deemed tolerable with it.  6-8 cm thick aluminum plate would be required.  That is quite out-of-line with current spacecraft hull design practices:  something near a millimeter.

Other materials based on polymers,  and even most rocket propellants,  are light enough molecules to be effective shielding with a low secondary shower risk,  yet with densities in the same ballpark as water,  for a thinner layer thickness.  Any of these could be practical shielding materials.

What you have to do is not simply add shielding weight to your design,  but instead rearrange the distribution of masses you already otherwise need,  so that they can also serve as radiation shielding.  You will need meteoroid shielding and thermal insulation,  and any manned craft will have water and wastewater on board,  as part of the life support system.  All spacecraft will need propellant for the next (and subsequent) burns.  You use a combination of these,  acting together.

The real suggestion here is to use water,  wastewater,  and next-burn propellant tankage as shadow shields,  in addition to the meteoroid protection and thermal insulation materials that the manned modules require anyway.  It doesn’t take much of this to cut the worst 60 REM/year GCR to under 50 REM/year.  It takes only a little more to cut worst case SFE to safe short-term exposure levels. 

If you cannot protect the whole manned interior,  then the flight control station becomes first priority,  so that maneuvers can be flown,  regardless of the solar weather.  Second priority would be the sleeping quarters,  to reduce round-the-clock GCR exposure further. 

See Figure A-3 for one possible way to do this,  in an orbit-to-orbit transport design concept.  This would also be a baton-spin vehicle for artificial gravity during the long transit.  Plus,  it requires a lot of interior space for the mental health of the crew.  Somewhere between 100 and 200 cubic meters per person is needed,  and it must be reconfigurable as desired by the crew.  That is a topic out of scope here.

Spin-up is likely by flywheels in the center module.  The vehicle is spun-up after departure,  and de-spun before arrival.  If a mid-course correction is needed,  the vehicle could be de-spun for that,  and spun back up for remainder of the transit.

Note how the arrival propellant and the water and wastewater tankage has been arranged around the manned core to provide extra shadow shielding for really effective radiation protection.  The manned core modules are presumed insulated by polymeric layers that also serve as meteor shielding (while adding to the radiation protection without being driven by that issue).  The pressure shell on the inside of this insulation should be unobstructed by mounted equipment,  so that easy and rapid access for patching is possible.

At departure,  the vehicle can be propelled by a different propellant and engine choice,  since departure is a short event.  The arrival propellant is likely a storable to prevent evaporation losses.  Return propellant tankage sets can be sent ahead unmanned,  for docking in orbit at the destination.  

There is an emergency return capsule (or capsules) mounted at the center module,  enough for the entire crew.  “Bailout” at destination presumes a rescue capability there.  Emergency bailout,  upon a failed burn for returning to Earth orbit,  is the main function of this capsule.  Routinely,  it could return a crew from the spaceship,  once parked safely in Earth orbit.

This kind of orbit-to-orbit transport design could serve to take men to Mars or to the near-Earth asteroids and comets.  For Mars,  the lander craft could be sent ahead unmanned to Mars orbit.  None are needed to visit asteroids.  The design of Mars landing craft is out-of-scope here.

By refueling and re-supplying in Earth orbit,  such a manned core design could easily be used for multiple missions,  once built.  Care must be taken in its design and material selection to support many thousands of cycles of use.  Thus the craft could safely serve for a century or more,  updated with better propellants and engines as the years go by.

There I went and wrote a basic “how-to” document for practical interplanetary spaceship design!

Figure A-3 – Using Otherwise-Required Materials To Also Serve As Radiation Shielding

Addendum B 8-19-19:  “Corporate Workfare” and Political Pork Versus Worthwhile Projects

The pork barrel aspect got started with the early days of NASA itself.  NASA labs and centers got situated in the districts of powerful senators and representatives,  in order to get the votes for the funding to carry out the Kennedy mandate to go to the moon.  Among many examples is the manned spaceflight center in Houston,  Texas.  This is essentially political corruption,  just never referred to by that particular word. 

Once this happened,  it became the norm,  and is seemingly not changeable.  A huge base of contractors and factory locations got developed to build the Saturns and the Apollo vehicles.  Once that program was done,  then in the minds of those congressional figures,  there needed to be something else for these contractors and factories to do,  because they were located in the districts of those powerful people.   Essentially the same contractor and factory base did the space shuttle,  and the ISS,  so the funding essentially continued to flow to those same districts,  for all these decades since. 

The current space launch system (SLS) giant rocket,  and the Apollo-on-steroids Orion spacecraft,  derive from that same contractor and factory base,  located in those same districts of those same powerful congressional seats.  Only the specific seat holders have changed.  The projects which NASA can take on have to use those assets in those districts,  lest funding dry up.  That is how congress dictates the details of what rockets and spacecraft get developed and what the flagship missions are.  NASA does not get to make a truly logical choice,  only the political choice.

SLS got started as “Constellation” fpr a return to the moon,  but is essentially the same rocket and capsule system,  just under multiple names and programs.  It is really just Saturn/Apollo redone with space shuttle technology and hardware.  This is now very old and inherently-expensive technology,  compared to what the new entrants (initially Spacex) in the business have been doing. 

In NASA’s early days,  the Mercury and Gemini capsules were supplied by McDonnell-Douglas,  and the Apollo capsule by North American Aviation.  All have since been gobbled-up into Boeing.  Rocket stages were supplied by Boeing,  General Dynamics,  McDonnell-Douglas,  and many others. 

Now there are just Boing and Lockheed-Martin (sometimes together as ULA),  and Orbital ATK,  plus the new entrant Spacex and perhaps soon Blue Origin.  All of the big-motor solid rocket plants (Thiokol,  Hercules,  CSD,  and others) are now part of Orbital ATK as an effective solid propellant monopoly. 

The maker of the Apollo lunar lander module (Grumman) is now part of Northrup-Grumman,  the third remaining main airframe provider,  and usually relegated to team member status,  in teams headed by either Boeing or Lockheed-Martin.  There is no longer any effective competition. 

In effect,  the many competing contractors in NASA’s early days have consolidated to a rather noncompetitive very few,  an oligopoly.  Without effective competition,  there is little incentive to actually go and be successful flying anything.  This shows in the track record of late:  it was 8 years from Kennedy’s mandate before any Saturns had flown at all,  to the first lunar landing using the final big Saturn 5.  Compare that to the totality that is SLS / Orion:  started under G. W. Bush,  continuing today,  a decade later,  billions over budget,  and still yet to fly at all.  This is about adapting existing engines;  the original Saturns were about developing new engines “from scratch”,  a much tougher job to do. 

Between the powerful pork-barrel aspect,  and the non-competitive oligopoly aspect,  it should be no surprise at all that what was once a powerful conglomeration of American know-how has devolved into little more than a welfare system for corporate giants sucking at the public tit,  getting their contract payments,  without really having to succeed at anything. 

Addendum C 8-19-19: Overall Mission Architecture and Vehicle Concepts For Mars

Fully covering this topic is way too large for an addendum to this posting.  It is essentially a new posting defining a planned mission to Mars,  and likely another separate posting describing how to use the same hardware to visit a near-Earth asteroid.  This 2019 Mars mission will be an update to an earlier posting titled “Mars Mission Outline 2016” and dated 28 May 2016.  Watch for these new postings.  They are in work as of this writing.

The new 2019 version of the Mars mission uses a larger orbit-to-orbit transport than the 2016 version,  and it also recovers the solar-electric tugs that preposition unmanned assets at Mars for the manned mission.  It uses similar landers as the 2016 version,  and it still jettisons the Earth departure stage without recovery. 

That non-recovered Earth departure stage could be addressed in future versions by fitting a larger departure stage with a second propulsion system,  possibly electric,  and putting it into a 2-year-period solar orbit after stage-off.  Then it could be captured into Earth orbit for reuse. 

Main point here:  if one does spin gravity in a baton-spin mode,  the resulting orbital transport vehicle is ill-adapted for a direct entry at Mars,  or a direct entry at Earth upon return.  Such a design is far better-adapted as an orbit-to-orbit transport,  with any Mars lander function relegated to a separate vehicle,  sent ahead separately with its propellant supplies.  Long-life reusability also points toward an orbit-to-orbit transport design,  free of entry heat shield requirements.  Such a concept was sketched in Figure A-3,  located in Addendum A above. 

This has inherently-higher velocity requirements,  there is no way around that!  But the direct-entry scenarios simply cannot provide this degree of safety for the crew in terms of radiation exposure,  microgravity diseases,  and having a “way out” at every step of the mission.  That higher velocity requirement is just the price you have to pay to do this job “right”,  in terms of ethical requirements!

The resulting mission architecture requires that both landers and the return propellant get sent ahead unmanned to parking orbit about Mars,  with the manned orbit-to-orbit transport arriving afterward,  and docking in Mars orbit with those items.  This powerful concept is not unlike the Lunar Orbit Rendezvous architecture that made it possible to mount each Apollo landing mission with only one Saturn 5 booster. The concept is illustrated in Figure C-1,  and its orbital velocity requirements in Figure C-2.  The mass ratio-effective velocity requirements are given in Figure C-3. 

For Mars arrival only,  there needs to be an additional propellant allowance to cover rendezvous requirements with the assets sent ahead.  As a wild guess,  add 0.2 km/s to the Mars arrival delta-vee requirement in Figure C-3.  There is no such corresponding allowance requirement for Mars departure,  and none is needed for the Earth arrival. 

 Figure C-1 – Overall Mars Mission Architecture Requiring Mars Orbit Rendezvous

 Figure C-2 – Orbital Velocity Requirements For The Orbit-to-Orbit Manned Transport

 Figure C-3 – Design Velocity Requirements For The Orbit-To-Orbit Manned Transport

The landers themselves are envisioned as one-stage reusable articles that make multiple flights,  based out of low Mars orbit.  Sending 3 landers allows one vehicle to make a landing,  with another in reserve as a rescue craft.  Thus,  there is a “way out” even during the landings,  unlike with Apollo.  The presence of a third lander allows one vehicle to become unserviceable,  while still maintaining the reserve rescue lander capability,  without which landings so far from Earth become too ethically risky to attempt. 

There is an insignificant velocity requirement for the deorbit burn,  with aero-deceleration to about Mach (1 km/s) at a rather low altitude.  From there the vehicle speed must be quickly killed with retro-propulsion (no time to deploy a chute,  much less wait for it to do any effective speed reduction).   That last requires a fairly large “jigger factor” to cover maneuvering,  hovering,  and diversion-away-from-hazards.  Even so,  the mass-ratio-effective velocity requirement for descent is not large at all.  This allows larger descent payload fractions.

The ascent velocity requirement must cover full orbital speed,  plus aero and gravity losses,  and a final rendezvous allowance.  This is a far larger velocity requirement,  but the payload is smaller because most of the supplies are exhausted,  some of the equipment will get left behind,  and the weight of the collected samples is not much in comparison.  The ascent payload fraction is quite a bit smaller.  These requirements are illustrated as a part of Figure C-4.  

Figure C-4 – Surface Landing Forays Based Out Of Low Mars Orbit

The unmanned transfers can be done more efficiently with solar electric propulsion (SEP) because of its far-higher specific impulse,  and because there is no need to transit the Van Allen belts quickly.  There is also no need to worry about reducing crew confinement times,  because these transfers are unmanned. 

This prepositioning of assets at Mars using SEP was also a part of my 2016 Mars mission posting.  The difference here is that I recover the SEP “tugs” into Earth orbit,  for reuse on future missions. 

The propellant sent ahead to Mars is all storable (the same NTO-MMH),  and comprises both the Mars departure and Earth arrival propellant,  plus a supply for the landers.  All this gets sent with the landers themselves,  as three unmanned “cluster” vehicles.  The Earth arrival propellant is arranged about the periphery of the manned spaces of the orbital transport,  so that it can also serve as part of the radiation shielding during the transit home. The Mars departure tankage is on one end,  and is staged-off after the departure burn,  before the vehicle is spun up.  See again Figure C-1. 

For the assets sent ahead with SEP,  mass ratio-effective design velocity requirements are much more problematic.  There are no drag losses in vacuum,  but the gravity losses are huge,  since the “burns” are months long.  For a rough rule-of-thumb estimate,  just use twice the values in Figure C-3. 

There is a lot of detailed work yet to do and to document for this version of the mission.  That will be covered in the new posting.  Watch for it.