Wednesday, April 3, 2019

Pivot-Wing Spaceplane Concept Feasibility


In an earlier posting,  I described a unique folding-wing spaceplane idea and explored its feasibility.  This is described in ref.1 (list at end of article).  The fundamental idea was to move the aerosurfaces into the wake while entering the atmosphere from orbit dead-broadside to simplify the aerodynamics and reduce the number of possible aeroheating failure modes.  By folding the aerosurfaces that way,  the dead-broadside forces that would rip the aerosurfaces off,  could be avoided.  Once subsonic,  these surfaces could more easily be deployed for an airplane-like landing.

This posting describes another way to accomplish the very same goal,  one that avoids the need for complicated fairings and entry-capable streamlining design for the folding-wing hinge joint. The folding butterfly (V) tail is not a problem,  and is retained,  being mounted to the dorsal surface already in the entry wake zone.  Instead of folding wings,  this revised concept uses a pivoting wing,  rather similar to the Russian “Baikal” missile booster seen at multiple recent airshows. 

This article presents nothing but a design concept feasibility analysis.  Only the gross overall dimensions and characteristics get determined.  I look at a ballpark weight statement,  best-estimate wing loading,  and  estimated entry gees and heating for that,  as well as an estimated landing speed.  Not much more.

Concept

Figure 1 depicts the vehicle design concept in cartoon form (all figures at end of article).  Figure 2 depicts how this concept might be operated in flying practice.  The craft is a small spaceplane launched with stowed wing using an appropriate two-stage rocket booster.  The entire delta-vee to low Earth orbit (LEO) comes from the booster rocket.  The spaceplane,  to be useful for real missions,  must arrive with significant maneuvering delta-vee (for plane changes,  transfer orbits,  rendezvous,  and the like,  plus including its final small de-orbit burn).

Figure 1 shows the vehicle to be a high-wing airplane with a butterfly tail and a non-circular cross section.  The main heat shield is located on a nearly (but not exactly) flat belly surface.  The figure says its radius of curvature should be about 1.5 times the length (or more) in order to reduce equilibrium stagnation surface temperatures enough to enable use of a low-density alumino-silicate ceramic material.  The actual radius turned out to be 1.55 times the length,  but could just as easily be twice the length.  The noncircular cross-section shape is similar to that of the old Mercury and Gemini capsules,  so that from orbit,  highly-emissive reradiating metal “backshell” surfaces can be used,  hidden from hypersonic scrubbing in the wake,  and with negligible plasma radiation heating effects,  at only 7-8 km/s entry speeds.  The same applies to the stowed wing,  but perhaps not the V-tail.

The cargo bay is near the middle,  with the center of gravity and the wing pivot,  so that changes in payload require minimal trim adjustments. Propellant tankage is disposed ahead and behind this cargo bay.  These can be simple tanks within the mold line,  with some insulation to protect them from the hot reradiating skin.  The propellants are room-temperature storables,  most likely monomethyl hydrazine (MMH),  and nitrogen tetroxide (NTO),  so that the same propellant supply serves both the main engine and the attitude thrusters,  with hypergolic ignition.  Simple is more reliable.

The cockpit is located forward,  and is the only pressurized space,  sized for a crew of two.  Since this craft returns to land while flying like an airplane,   having a second pilot to support the first serves safety and reliability well,  just as in airliner flying.  It is likely this craft will be a challenge to fly,  as the return trajectory in Figure 2 suggests.  That makes two pilots more desirable in any event.

Anticipating high landing speeds because of the geometric limits for the size of the pivot wing,  landings on dry lake beds are presumed,  which makes a landing gear arrangement like that of the X-15 desirable and proven.    There is a steerable nosewheel forward under the pressurized cockpit,  and a pair of main skids near the tail in the engine compartment.

Pitch,  yaw,  and roll are presumed controlled with the V-tail surfaces,  which otherwise also hinge so as to stow vertical in the wake for entry,  and afterward deploy to about 45 degrees off vertical,  for aerodynamic flight.  Pitching tail surface leading edges together up-and-down provides pitch control,  pitching them opposite provides forces that affect both yaw and roll,  requiring different amounts from each fin to properly allocate the yaw and roll effects,  as influenced by the high wing.  The wing can be a subsonic airfoil,  and is a straight wing of fixed geometry,  effectively high-wing-mounted.

Unlike the space shuttle,  but like the other folding wing concept in ref. 1,  this craft enters the atmosphere at essentially 90 degrees angle-of-attack (AOA) and zero roll angle.  The flat shape of its belly provides significant lift with small changes in pitch,  so that aerodynamic lift can be used to fly the desired entry trajectory (a technique well-proven with Apollo and Space Shuttle).  The intent here is to return with little-to-zero payload and near-zero propellants,  having only a small allotment for the attitude thrusters. The heavier the return payload,  the higher the landing speed. 

The craft comes out of hypersonics at about Mach 3 at very high altitude (near or above 100,000 feet) still dead-broadside to the airstream.  Closer to Mach 1,  the V-tail and attitude thrusters put the nose down streamline,  and a drogue chute deploys,  that is sized for about the same drag as that of the entry configuration dead-broadside (both are sufficient to reach subsonic terminal velocities).  Once streamwise and subsonic,  the pivot wing is deployed and the drogue chute discarded.  At this point,  the craft becomes a straight-wing V-tail glider,  handling very much like any subsonic airplane,  just flown dead-stick (with maybe just enough propellant still on board,  to support a go-around on the main rocket engine).

Sizing

Figure 3 depicts the spreadsheet worksheet used to rough-out the basic weight statement and characteristics of the design,  as a function of overall length and selected “wing loading” values.  User inputs are highlighted yellow.  Most (but not all) significant outputs are highlighted blue or green. 

The first data group is “engine”  and gives delta-vee capability for given specific impulse (Isp) and the mass ratio values that come from the weight statement.   The engine and thrusters should do as well or better than the 300 sec of Isp shown,  and the mass ratio-derived ideal delta-vee value exceeding 2.5 km/s at 300 sec is quite attractive for a variety of possible missions.  One must hold in reserve at least the deorbit burn and an allotment for the attitude thrusters.

The second data group is “inert weight fraction”,  and is just an organized way to guess a realistic inert weight fraction,  based crudely upon what the structure must do.  These methods are described more in ref. 2,   as part of a larger methodology for estimating performance of rocket stages. The result here of 20% should be quite realistic.  Bear in mind that operational military and commercial airplanes here on Earth usually run near 40% inert,  where that category plus propellant fraction,  plus payload fraction,  must sum to 100%.

The third data group is “payload”,  and shows 200 kg for two men,  a quarter ton for their suits and life support,  and 5 tons max in the cargo bay.   The user inputs a payload fraction (in this case 20%),  and the remainder is the propellant load.  That leads immediately to the weight at ignition,  and thus the vehicle weight statement in the fourth data group. 

This spreadsheet analysis simply presumes that the body planform area is 0.8*length,*width,  and that the body cross section area is 0.8*width*height.  It also presumes the chord of the pivot wing is 1/3 the body with,  and that the span of the pivot wing is ¾ the body length.  That leads to a fixed wing area to body planform area ratio of 31.25% or thereabouts.  The user inputs the ratios of body width/length and body height/length,  representing fineness ratio proportions (both 16%,  or 6:1,  here).

There is a user input for the cargo bay length/body length proportion,  that eventually leads to a cargo specific gravity,  under the assumption that cargo fills 100% of the available volume.  I set that for a specific gravity 1/3 that of water,  to represent bulky,  lower-density items.

In the “aerosurfaces” group,  one sets the tail proportion and the entry “wing loading” of burnout weight/body planform area,  along with a hypersonic drag coefficient for the intended shape,  yielding an entry ballistic coefficient.  That gets used in the entry ballistic analysis.

A representative max wing loading for airplane-like flight with the wing deployed,  would be burnout weight divided by the sum of body planform area plus wing planform area.  That is because,  while the body lift curve slope is low,  the body planform is the larger area,  and thus a significant contributor to lift.  This applies,  as a user-input max lift coefficient,  to the landing speed calculation group.

The “proportions” group is where one inputs the body length,  its width and height ratios,  and the cargo bay length fraction.  This is where the various areas and volumes get estimated,  along with the cargo specific gravity,  and the weight/area loadings. 

It is necessary to iterate to closure here.  The weight/area outputs from “proportions” must match those derived from your input weight/area loading in “aerosurfaces”.  You have “body length L, m” and entry loading “Wbo/Abdy,  psf” as your values to change until you achieve convergence.  The higher the Wbo/Abdy figure,  the higher the ballistic coefficient will be,  and the higher the landing speed will be. 

I found that guessing max lift coefficient for landing was too unreliable.  So I added a worksheet to estimate this more explicitly from the “proportions” outputs.  This is the “landing” worksheet,  shown in Figure 5.  That worksheet produces the “right” stall lift coefficient to use in the “landing” group of the “rough-out” worksheet (and then you will see the landing speed estimates agree between the two worksheets).  I also added a worksheet to estimate the size of the drogue chute,  shown in Figure 4.

The landing worksheet estimates lift curve slope for the very low aspect ratio “wing” that is the body,  from an equation obtained from ref. 3,  the Hoerner – Borst lift book that is analogous to Hoerner’s “drag bible” (ref. 4).  This would be equation 9,  located on page 17-3 of chapter 17 in Hoerner and Borst (ref. 3).  Low aspect ratio wings inherently have very low values of lift curve slope. 

The only additional inputs to the “drogue” worksheet,  beyond outputs from “proportions” in the “rough-out” worksheet,  are the parachute subsonic drag coefficient and the end-of-hypersonic (Mach 3) point from the entry trajectory analysis.  The drogue is sized to provide the same drag and subsonic terminal speeds at 60,000 feet and 20,000 feet altitudes,  as the body falling dead broadside with the wing stowed.  The end-of-hypersonics point is just a check:  need 100,000 feet (30 km) or higher.

Entry Analysis

The entry trajectory analysis is a very simplified 2-D Cartesian model from the mid-1950’s that was used for warhead entry analysis.  It is based on a scale-height model of approximating density versus altitude,  and presumes a constant trajectory angle in 2-D Cartesian space.  To use it for estimates here requires that one fly a trajectory always oriented at a constant angle to local horizontal around the Earth. The range wraps around the Earth. The analysis is attributed to H. Julian Allen,  and is described in ref. 5.

In my spreadsheet version of the old model (image given in Figure 6),  there are user inputs for the vehicle model,  the scale height model,  the entry interface conditions,  and the stagnation heating model.  The vehicle model requires a ballistic coefficient and a “nose” radius (really the heat shield radius of curvature).  The entry interface model is altitude (for LEO,  140 km),  velocity (for a surface-grazing ellipse,  7.742 km/s),  and path angle below horizontal (for that same surface-grazing ellipse,  2.35 degrees) at entry interface conditions.  The final vehicle model achieved here has a ballistic coefficient of 439 kg/sq.m,  a length of 17.45 m,  and a heat shield radius of curvature of 27 m. 

Use of this spreadsheet model requires inserting a row of cells to represent the altitude and results for a speed corresponding to Mach 3 end-of-hypersonics (in this case about 1 km/s).  One iteratively adjusts the altitude so that a 1 km/s speed shows in the table.  One uses data from start (at entry interface) to only end-of-hypersonics for creating plots.  The model does not apply once speed is no longer hypersonic.  That is why you stop at the Mach 3 point for bluff bodies.  All of this is shown in Figure 6.

As also shown in Figure 6,  I added two things at the bottom.  One picks off the metric-units peak heating rate (wherever it occurs),  and the integral of heating at end-of-hypersonics,  and inputs them to a US customary units converter.  Next to this,  one uses an input emissivity and the converted peak heating rate to estimate the surface temperature at peak heating,  under the assumption that reradiated cooling power equals the convective heating power.  This would apply to a refractory (non-ablative,  and not-liquid-cooled) heat shield. I converted to US customary,  because that is the units of the radiation constant that I know,  and those are the materials-limitation properties that I know.

It is easy enough to highlight where the instantaneous gees exceed 5,  determine the peak gees,  and use the time scale to estimate how long the high-gee interval is,  that must be endured.  It is also easy to determine whether hypersonics is over at (or above) 100,000 feet (about 30 km),  as it should be for the rest of the concept’s descent sequence.

It is easy to plot the data from the entry spreadsheet analysis.  These are given in Figures 7-10.  Figure 7 is a range versus altitude plot,  illustrating the constant angle trajectory in the 2-D Cartesian model.  Both slant range down the trajectory,  and horizontal range along the ground,  are shown.  At only 2.35 degrees different,  the two curves fall on top of each other in this plot.  Horizontal range wraps around the curvature of the real Earth,  and the constant descent angle must be treated as constant with respect to local horizontal as one proceeds along the trajectory.

Figure 8 shows velocity versus altitude.  It starts at 140 km altitude and orbital speed,  and ends just under 35 km at 1 km/s (just about Mach 3).  Not much deceleration happens at all,  until one descends to about 60-70 km.  From there deceleration quickly grows to high values at about 40-50 km and below.

The two key kinematics results are shown in Figure 9.  These are velocity versus time,  and deceleration gees versus time.  Peak gees is about 6.22,  at 326.9 sec,  where the velocity is 3.820 km/s at 42.5 km/s altitude.  The time above 5 gees is only about 30-40 sec.  The peak and duration of the high-gee exposure is feasible for a seated astronaut,  to be endured in the “eyeballs-down” direction.   

Figure 10 gives the time history of the convective stagnation heating rate as q, W/sq.cm,  and its time-integral accumulation of energy Q, KJ/sq.cm.  Peak heating rate occurs a little earlier than peak deceleration gees,  being 26.75 W/sq.cm at time from entry interface 270.8 seconds,  altitude 55 km,  and velocity 6.824 km/s. 

End of hypersonics (at just about Mach 3) occurs at 412.6 seconds from entry interface,  altitude 34.78 km,  and speed 0.999 km/s. Looking at the heating rates,  a good guess says the plasma-induced radio blackout is about 3 minutes long,  as expected.  The whole entry is a bit over 6 minutes from interface to end-of-hypersonics.  These numbers are very,  very realistic,  despite the oversimplified analysis method.  It looks like my misuse of the old warhead entry analysis is justified for capsule-like entry.

Feasibility

The first time through,  I used a shorter (13.5 m) vehicle with a higher ballistic coefficient (732 kg/sq.m) and a 27 m heat shield radius,  which had an infeasibly-high max-load landing speed near 300 mph,  and came out of hypersonics at about 31 km.   It showed a peak surface temperature of 2541 F,  too high for an alumino-silicate refractory heat shield material (shrinkage cracks form above 2350 F upon cooldown). It was at this point that I added the drogue and landing worksheets to the rough-out worksheet,  in order to better optimize this design concept.

The final form is a 17.45 m long craft,  with a lower ballistic coefficient of 439 kg/sq.m,  and the same 27 m heat shield radius of curvature.  That reduced peak gees and peak heating,  reduced the heat shield temperature to a barely-feasible 2345 F,  raised the end-of-hypersonics to nearly 35 km,  and lowered the max-load landing speed to about 217 mph at sea level stall (under 200 is desired). 

These were computed for the full burnout weight loaded onto the body planform or total planform areas,  meaning flying back with full cargo.  Flying back with reduced cargo will lower heat shield temperature and landing speed.  That improves the feasibility of this roughed-out design.

Having a stagnation-point surface temperature under 2350 F is very important if one wishes to use a low-density alumino-silicate ceramic as a refractory,  re-radiation-cooled heat shield.  This need not be the logistical nightmare that Space Shuttle tiles proved to be.  There are other materials that could be developed with the applicable characteristics,  and providing the redundant retention that shuttle tiles lack.   See Ref. 6 for a very experimental material that was a fabric-reinforced low-density ceramic. 

Conclusions

What this analysis shows,  very much like that in ref. 1,  is that this sort of small spaceplane is within the realm of engineering feasibility.  The pivot-wing design would be easier to implement as entry heat-protected than the folding-wing design of ref. 1.  All-in-all,  borrowing the Russian “Baikal” pivot-wing approach is an improvement,  provided that it is deployed subsonically to reduce aerodynamic deployment loads.  It is limited in how much wing area can be feasibly added in a dorsal-only mount. 

The craft as-sized is 28.5 metric tons at fully-fueled,  fully-loaded ignition.  Its body is 17.45 m long,  and about 2.8 m wide,  and 2.8 m high.  Only the tail fins stick out to the dorsal side.  It might actually fit within the standard payload shroud of a Falcon-Heavy booster rocket,  and certainly falls within the payload weight limit for that rocket to recover its first stage cores.   If SLS ever really flies,  it could certainly carry one (or more) of these craft.

About 2.4 km/s worth of on-orbit delta-vee makes a great many missions possible with a craft like this,  once delivered to eastward LEO by a suitable booster.  That is over 15 degrees worth of plane change,  or very nearly to Earth escape velocity.  Multiple orbit visit locations in one mission become possible,  a very attractive characteristic indeed.

Having a small airplane with an easily-stowed wing as the returning spacecraft,  makes possible picking this up with something like a C-130,  and flying it to any suitable launch site.  Having a low-density alumino-silicate heat shield makes a long service life between repairs feasible,  as long as it does not take the form of bonded tiles,  no two of which are alike,  as with the Space Shuttle.  Thus logistics are greatly simplified.  That makes turnaround time shorter,  and flying costs lower. 

Final Comments

This is not a real design study.  It is only a configuration rough-out and basic feasibility analysis.  It shows that such a design really is feasible,  that much is certain.  Little else.

But,  the reader is cautioned to not take this work to be more than it actually is!  I ran no dimensions other than some overall ones,  selected no materials,  conceived and weighed no structural components,  and I did not do any detailed heat transfer,  air loads,  or stress-strain analysis.  Most of the parts for which those kinds of design analyses are appropriate,  have not been designed at all.

This pivot-wing design approach offers a much easier-to-heat-protect method of mounting the stowable wing for reentry.  It is strongly limited in how large that wing can be,  relative to the rest of the airframe.  Thus it inherently suffers from high landing speeds.

The earlier folding-wing concept can have a much larger wing relative to the rest of the airframe,  which means it can have a much lower landing speed.  That offers any airport as a landing field,  even if the main skid gear is retained (wheels add more weight).  The problem is heat-protecting the hinge joint,  especially if a low-wing design.  That is not impossible,  just quite difficult.

References

#1.  Johnson,  G. W.,  “A Unique Folding-Wing Spaceplane Concept”,  article posted on http://exrocketman.blogspot.com,  dated March 2,  2013.

#2. Johnson,  G. W.,  “Back-of-the-Envelope Rocket Propulsion Analysis”,  article posted on http://exrocketman.blogspot.com,  dated August 23,  2018.

#3. Hoerner,  S. F.,  and Borst,  H. V.,  “Fluid Dynamic Lift”,  published by Mrs. Liselotte A. Hoerner,  1975.

#4. Hoerner,  S. F.,  “Fluid Dynamic Drag”,  self-published by the author,  1965. 

#5. Johnson,  G. W.,  “BOE Entry Model User’s Guide”,  article posted on http://exrocketman.blogspot.com,  dated January 21,  2013.

#6.  Johnson,  G. W.,  “Low-Density Non-Ablative Ceramic Heat Shields”,  article posted on http://exrocketman.blogspot.com,  dated March 18,  2013.


 Figure 1 – Vehicle Concept for Pivot-Wing Spaceplane



 Figure 2 – Operations Concept for Pivot-Wing Spaceplane



 Figure 3 – Image of Spreadsheet Worksheet Used For Vehicle Rough-Out Calculations

 Figure 4 – Image of Spreadsheet Worksheet Used to Size the Drogue


 Figure 5 – Image of Spreadsheet Worksheet Used for Landing Speed Calculations



 Figure 6 – Image of “BOE Entry” Spreadsheet Worksheet Used for Entry Estimates



 Figure 7 – Spreadsheet-Generated Plot of Entry Trajectory Shape



 Figure 8 -- Spreadsheet-Generated Plot of Entry Trajectory Deceleration Trend



 Figure 9 -- Spreadsheet-Generated Plot of Entry Trajectory Kinematics



Figure 10 --  Spreadsheet-Generated Plot of Entry Trajectory Heating


Monday, April 1, 2019

Something Added to the House


We recently got a historical marker on our house.

The first photo is of our front porch,  near the front door.  You can see the big metal Texas star,  which matches one on the farm shop across the driveway.  The new historical plaque is on the front wall just left of the metal Texas star.





















The second photo is a closeup of that historical marker,  so that you can actually read it.  Enjoy. 





















The main takeaway here is that historical and hysterical are spelled (and pronounced) very similarly.

Saturday, March 16, 2019

THIS is a Slide Rule!


This simple device is what equated to a modern scientific pocket calculator when I first entered the aerospace defense workforce (see the photo).  This is a slide rule,  the "calculator" used by engineers and scientists for 300 years before there were any electronic calculators at all!  This very slide rule is what I designed my first airplane with,  and my first half dozen supersonic missile propulsion systems with. 




For problems you could not handle pencil-and-paper with a slide rule,  there was the mainframe computer.  These were devices that filled a room the size of a small house,  air-conditioned to about 60-65 F,   so that the magnetic iron cores and wiring would not try to melt down!  You loaded your data and programs into the mainframe "card batch" in trays,  up to 2000 cards at a time,  using paper punch card technology.  Job turnaround time was measured in hours,  sometimes days. 

As for modern spreadsheet technology,  when I first entered the workforce,  repetitive calculations were manually laid out on a big piece of paper in a matrix format.  You ran the actual calculations yourself,  using a slide rule,  or a bit later,  a hand-held electronic calculator.  You filled in the matrix slowly,  literally doing each and every calculation yourself,  and finding out "up-close-and-personal" what could go wrong with the processing of the data.  That is where today's spreadsheet software came from!

For a given manual spreadsheet job,  this experience taught you exactly how to program the calculations into a scientific programming language (in those days,  something like an early FORTRAN or BASIC),  complete with all the processing logic and error-trapping.   That required punching the program statements onto cards,  for card-batch load and debugging (again,  job turnaround time was hours-to-days between each run).  Once you did this,  you could do similar jobs,  requiring the exact same analysis,  far more quickly.  (Or you could modify your program to handle other jobs that were similar,  but slightly different in a few details.)

Most people today do not realize this,  but NASA mission control in Houston did not have any real computer consoles until the Space Shuttle first flew!  During the earlier Mercury,  Gemini,  and Apollo programs,  those flight controller consoles were only keyboard-controlled communication displays,  each slaved to a counterpart in a back room outside the mission control room.  There was a team of people (both men and women!) in each back room,  who answered the flight controller's question with slide rule calculations,  and typed in the answers,  so that their numbers appeared on-screen in mission control.  That,  plus analog instrument readouts converted to digital format on-screen,  is literally all the mission flight controllers had to work with!

With the exception of a mainframe-computed figure-eight orbit between Earth and moon,  NASA literally sent men to the moon during Apollo with slide rules (just like the one in the photo)!  And,  the record-breaking X-15 rocket plane (and all its earlier progenitors),  plus the SR-71 jet aircraft,  and all the early supersonic jet fighters,  were designed with nothing but slide rules.  Same for all their rocket and gas turbine engines!  And their heat protection schemes.

My slide rule still works.  I still use it when the electronics conk out,  which they inevitably do,  occasionally.  The slide rule never conks out like that.  It's just slower. 

Monday, February 4, 2019

Designing Rough-Field Capability into the Spacex Starship

Update 2-5-19 is at the very end,  after the article figures.

----------

Bear in mind that Starship is Spacex’s new name for what was once the BFS second stage spacecraft article of its BFR/BFS system.  To be useful at Mars or the moon,  this spacecraft must be able to make rough-field landings.  Its mass is heaviest when fueled for launch.  Initially,  it must be refueled for use at Mars,  and eventually,  also the moon.  These static loads are larger than the landing weights,  even factored for dynamical impact.

There are two parts to this:  (1) tip-over on rough or sloping surfaces,  and (2) not exceeding the bearing load capability of the natural surfaces.  There is also a new idea presented here for creating very large landing pad surfaces that fold so as not obstruct airflow,  in a very practical way. 

The tip-over problem was well-explored in another article on this site,  as part of updates to the basic performance evaluation article.  That article was “Reverse-Engineering the 2017 Version of the Spacex BFR” dated April 17,  2018.  The same article identified soil bearing strengths as likely inadequate to support the spacecraft when refueled for launch. 

The related article “Relevant Data for the 2018 BFS Second Stage” dated September 24,  2018,  included among other things a way to reconfigure the round tip-mounted landing pads into oblong pads of increased area.  Those results were still inadequate for the loose fine sand-like surfaces of much of Mars. 

What is analyzed here is a different landing pad idea,  depicted in Figure 1 (all figures at end of article).  Essentially,  panels resembling landing gear bay doors are built into each side of each fin tip,  with hinge lines at the aft trailing edge (which is the touchdown surface otherwise). Unfolded hydraulically,  these panels become very large landing pad surfaces.  Folded,  they do not protrude into the ascent or descent airstreams at all. 

The same figure shows an 1100 metric ton fueled mass,  which really could be as large as 1300-something tons.  However,  the BFS weight statement is still not known publicly with any certainty.  This figure is in the same ballpark,  given all the other uncertainties. 

Assuming 1100 metric tons of mass,  the weights on Earth,  moon,  and Mars are given in the figure.  The bearing pressures associated with those weights are also shown,  assuming 2 m by 2 m parallelogram-shaped folding panels are used.  This assumes 2 panels per fin,  and 3 fins.  Just that initial assumption provides some 10 times the bearing area as the roughly 1 m diameter round tip pads shown in the Spacex illustrations of this vehicle. 

Figure 2 presents safe load-bearing strengths for civil engineering purposes of various types of Earthly surfaces.  These came from an older-vintage Marks’ Mechanical Engineer’s Handbook.  No such reference yet exists for lunar or Martian soils.  However,  experiences from the Apollo missions verify that the lunar regolith is similar to fine,  loose Earthly sand. 

Experiences with the various Mars landers and rovers suggest that much of Mars is similar to lunar regolith and to Earthly fine,  loose sand.  Some of Mars seems to have a mix of sand,  gravel and larger rocks,  perhaps similar to Earthly surfaces such as loose beds of medium and coarse sand,  or perhaps even as substantial as beds of coarse sand with gravel.  These require picking,  not a spade,  to remove.    A hard clay requiring picking would also be of similar bearing strength. 

All of this is indicated in Figure 2.

Figure 3 compares applied bearing loads to soil strengths for the moon,  Mars,  and Earth.  The moon is the least demanding problem because of its lowest gravity.  A design adequate for the moon is adequate only for some of Mars:  the indicated folding-panel parallelogram dimensions would be 2.2 m by 2.2 m,  for about 29 sq. m bearing area.  On Earth,  sites must be strong,  well-packed sand/gravel or hard clay.  No beaches,  no sand dunes,  no soft desert.

If we increase that pad area to be capable all over Mars,  the folding panel parallelogram dimensions become 2.67 m b 2.67 m,  for about 43 sq.m bearing area.  This is part of the comparative pad area summary and comparison given in Figure 4.  There is a trade-off here:  the bigger these folding panels can be,  the more of Mars (that is otherwise fairly smooth and level) is a feasible landing and takeoff site.  The moon is not a problem,  nor is much of dry land Earth (most anything requiring picking or blasting would be adequate). 

This would simply not be the case with those 1 m diameter round tip-mounted landing pads that Spacex illustrates.  That total pad area is about 2.4 sq.m,  give or take a small amount.  What is needed for rough-field capability on Mars (or the Earth) apparently falls in the 30-45 sq.m range.  You simply cannot do that,  in any practical way,  with smallish fixed-geometry tip-mounted landing pads.  Those will require thick reinforced-concrete landing fields,  or else thick solid rock.

I know they have their hands full at Spacex trying to make this vehicle a reality.  But some of the thinking I have explored here,  also needs to go into Spacex’s designs! 


 Figure 1 – Fold-Out Panels as Rough Field Large Landing Pads

 Figure 2 – Representative Soil Bearing Strengths for Various Earthly Surfaces

Figure 3 – Relating Landing Pad Areas to Soil Bearing Strengths


Figure 4 – Tradeoff of Increasing Pad Size vs Landing Site Choices



Update 2-5-19:

As a follow-up,  I put some more-traceable masses for the BFS/”Starship” weight statements into a spreadsheet,  with gravity data,  and empirical data for the safe bearing capability of various Earthly surfaces.  This included the selections (and rationales) as for which surfaces resemble the moon and Mars,  and which might be the widespread worst cases for emergency landings on Earth.  Sources are indicated.

Masses and gravity are given in Figure A below (all figures at end of this update). The residual propellant remaining is nothing but a wild guess,  knowing that a dry-tanks landing is truly risky.  As it turns out,  the presence or absence of residual propellant mass at touchdown does not drive the sizing of landing pad area.  Refilled launch weight drives this. 

The safe surface bearing-pressure capability data for a variety of Earthly surfaces is given in Figure B.  These are civil engineering data from an old-vintage Marks’ Mechanical Engineer’s Handbook.  These surfaces are rather variable in properties,  as is typical of geology.  They represent safe bearing pressure loads so that your structure or object does not try to sink slowly into the surface,  even over long periods of time.  It is conservative,  ethical practice to use the min pressure values for design.  In the handbook,  the table presented both metric and US customary values,  for which it was obvious the metric were converted from the US customary source values.

To this table I have added the notations about which surfaces resemble the bulk of the moon and Mars,  and the rationales for those selections.  I have also indicated the most common soft-surface emergency landing surfaces for Earth,  excluding soft sand beaches and deserts (and swamps).  The rationale for that is simple experience.  These are my best estimates,  if I had to do this.

I worked out local-weight weight statements for Earth,  Mars,  and the moon,  for two configurations.  One is arrival,  with the larger payload sent from Earth,  and only residual propellant left at touchdown.  The other is at departure,  with full propellant load,  but a reduced payload,  for the return to Earth.  This reflects exactly what Spacex says the scenarios are for Mars. 

Spacex says there needs to be no lunar refilling for return to the Earth from the moon,  but this analysis anticipates that eventually lunar refilling might be attempted,  for trips from the moon to destinations other than Earth.  It does not matter;  as it turns out,  the lunar launch case does not determine the design requirement for landing pad area.  But I had to check,  as a due diligence item.

What I found was that local-weight launch weights exceeded local-weight touchdown weights by roughly a factor of 5.  Therefore,  it is only local launch weights that govern max bearing pressure exerted upon the local surfaces.  One of these is the worst case that drives the design requirements.

Surfaces on the moon,  and the great majority of Mars,  resemble soft fine sand.  There are places on Mars with somewhat-stronger surfaces,  but these are definitely not the majority of possible landing sites.  You do not want a rough-field landing design restricted to rare site opportunities.  That would be rather pointless.  As for the range of properties,  you have to select the min capability. 

Figure C combines the local-weight weight statements with the appropriate selected surface bearing capabilities,  to produce min total landing pad areas for Earth,  Mars,  and the moon.  Mars governs,  and by a significant margin.  The total pad area result obtained here is not at all far from the seat-of-the-pants 43 sq.m in the original article just above.  But this updated result is more traceable,  and therefore the more reliable value.  It is just about 46.2 sq.m.

The same figure also gives a selection of parallelogram dimensions for each fold-out landing pad panel.  This is a function of the number of landing-leg fins,  the number of panels per fin,  and the aspect ratio of those panels (the height to base ratio for the parallelogram shape).  Beyond scope here is the structural design of such panels;  it seems likely that the lightest version would be nearer aspect ratio 1.  That would be a panel 2.77 x 2.77 m size,  vs 2.67 x 2.67 in the original article.  That’s pretty close!

Again,  I must point out that this is the sort of design,  and design analysis,  that is needed by Spacex to really provide a rough-field capability for its BFS/”Starship” spacecraft on Mars (or anywhere else).  The round tip pads Spacex currently shows are roughly a meter diameter,  for a total of about 2.4 sq.m total area,  roughly some factor 20 smaller than what I determined here. 

                Implications

Without the large total landing pad area I found,  the craft is restricted to very thick reinforced concrete landing pads,  or smooth,  level stretches of thick,  solid rock.  Otherwise,  while you might possibly land (and no guarantees about that!),  you’ll sink-in unevenly,  and tip-over “for sure”,  upon refilling propellant for launch.

At 2.4 sq.m and Mars launch weight,  the applied bearing pressure is ~0.39 MPa,  similar to the 0.38 MPa min capability of Earthly coarse sand & gravel,  and similar to only a minority of Mars!  And even that still lacks the factor 2 margin you need for the dynamical impact effects of the touchdown:  the fin tips will inevitably stab deeply into the surface,  risking getting stuck like tent stakes.  Very likely,  they will stab-in unevenly,  risking a tip-over,  even if the site is level.  So the chances of a successful landing with the depicted design are very poor,  and there is no chance at all of a successful refilled launch.

If the site actually resembles the soft sand that is the bulk of Mars,  there is no chance at all of a successful landing,  and by far.  The applied bearing load is 0.39 MPa,  even without the factor 2 for the dynamics.  The soft sand capability is only 0.1 MPa.  That’s about factor 4 outside-of-the-ballpark,  and about factor 8 wrong with dynamics allowed-for!  That’s a crash,  period.  That’s exactly what will happen with three 1-meter-diameter fixed-geometry fin-tip landing pads!

                Conclusion

Something about the design Spacex currently shows,  must change substantially,  before even an unmanned cargo ship flies to Mars.  They need 47 sq.m of pad area,  not 2.4 sq.m.  That’s what the best-available data actually says. 


 Figure A – Masses and Local Gravity for the Local-Weight Weight Statements




 Figure B – Surface Bearing Capabilities on Earth,  Annotated for Mars and the Moon



Figure C – Local-Weight Weight Statements and Bearing Loads Size Landing Pad Area


Thursday, January 17, 2019

Border “Crisis”? Nope

Update 5-5-19The continuing crisis has morphed somewhat.  There now seems to be a bigger population of migrants classifiable as refugees seeking asylum than guest workers looking for work. 

This is for reasons beyond our control:  the effect of at least 3 failed states in Central America.  This might be temporary,  or it might be permanent.  Statements by the various border control agencies do confirm this assessment,  whether worded that way or not.

The problem:  we are set up to pursue and deal with illegal migrant workers (a problem we created with artificially-low quotas for guest worker visas).  Most of these folks are Mexican residents,  not Central American refugees. 

We are not set up to deal humanely with refugee families from Central America.  THAT is what the border control agencies are really telling us they cannot deal with. 

I have already described what to do about the Mexican guest worker problem:  revise the visa quotas to reflect the demonstrable size of that particular job market. That is something for Congress to do,  and they have ignored this for ~7 decades now,  thus creating the 10-12 million strong illegal alien population in the US. 

My fellow citizens:  PLEASE HOLD YOUR CONGRESSIONAL REPRESENTATION ACCOUNTABLE FOR THAT FAILURE!  That applies to both the House and the Senate. 

If your current representation does not talk about fixing this,  then elect someone new.  You could NOT POSSIBLY do any worse!

I have also described the ONLY ethical way to handle the exploding refugee problem:  staff up and just deal with it!  More judges,  more personnel in general,  and more facilities appropriate to the task (and tent cities along the Texas and Arizona borders are NOT what I mean !!!). 

Several administrations in recent years have neglected this issue,  but the current Trump administration is quite egregiously failing to deal with this at all.  They are motivated by a self-evident "we want no brown-skinned immigrants at all" policy,  if you deign to call that a policy.

I do not:  this is racism far beyond the abhorrent "Jim Crow" type,   this is closer to that utter abomination demonstrated by the Nazi Germans during World War Two.

It manifests as who showed up as a lot of the pro-Trump supporters at Charlottesville,  VA,  for one. These included a lot of people wearing Nazi uniforms,  carrying torches and swastika flags,  and other Nazi regalia.  Despite what Trump said,  those are NOT "fine people". 

Face the truth:  there are no good Nazis!  There never were.

My fellow citizens:  PLEASE HOLD THE TRUMP ADMINISTRATION ACCOUNTABLE FOR THIS UTTER FAILURE TO CONFORM TO MINIMAL STANDARDS FOR HUMAN DECENCY !!!!   

I don't care whether this accounting is his non-re-election in 2020,  or his impeachment and conviction / removal from office sooner. 

Just do it!  Before he finishes destroying our country by artificially dividing it (with lies) for political gain (his playbook since entering the race in 2015).  His fact-checking rate is almost exactly 0%.

As for details about how to properly handle guest workers and refugees seeking asylum,  read on. I also deal with border smugglers (drugs or people). 

Update 2-3-19:  see below at end.
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This “crisis” is not about compromising on wall funding vs DACA,  it is about the hurtful practice of holding America hostage,  by means of damaging government shutdowns,  to get a political desire not otherwise obtainable.  This evil practice has to stop,  and this time around is as good a time as any,  to put a permanent stop to such behavior. 

The border “crisis” itself goes far beyond just walls and DACA.  “They” are lying to you when they cast it only in those terms!  Both sides in Congress,  and the administration,  chronically lie about this issue,  but the Trump administration has been (by far) the most egregious with its lying. 

Most of the so-called “news” about this is also a lie,  even if only lying-by-omission.  Be careful of your sources:  if you hear no divergent voices to your own opinions,  then you are in an echo chamber being fed propaganda.  “Propaganda” is just a long-winded way of spelling “lie”.

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Asylum seekers,  guest workers,  and cross-border smugglers are different problems with different solutions.  Only false political arguments lump them together.

Government statistics prove that asylum seekers and guest workers are less likely to commit crimes than US citizens at large.  There is no real threat there,  despite all the “justification” claims by the Trump administration.

“Less likely” is not zero,  there are bad apples in any barrel.  Finding the bad ones at the border crossings merely requires adequate staffing to do the job.  A wall does not help that.

Cross-border smugglers are the drug cartel and gang members;  those are the real threat.

Asylum seekers

The number of refugees seeking asylum at our southern border is up in recent years.  The reasons why are not under our control. 

These people have the legal right to cross the border and ask an official for asylum.  There is no question about that aspect of federal law,  despite all the political denials. 

This legal right was illegally denied by the Trump administration,  by means of “criminalizing” any crossing not at a port-of-entry,  then claiming that “criminality” as an excuse not to hear the cases and separate children.  That law is settled:  it does not care where the asylum-seeker crossed. 

Trump administration officials have admitted in public that they intended to use the threat of separation of children as a deterrent to stop other asylum seekers.  This is not just immoral,  it is evil.

It is a fact that we have too few immigration judges to hear these asylum requests.  The backlog is unconscionably high,  and getting far worse with the shutdown.

Our holding facilities were designed to handle men,  not women and children.  This plus the backlog leads to cages and tent cities.  These are an immoral evil,  that no one can deny.

A whole-border wall “fixes” none of this.

Why not repair the asylum process and staff-up to do it right?  That would be far cheaper than any wall,  and it frees up many of our agents to chase the cross-border smugglers instead!

Guest workers

This is a very old problem,  the result of about 7 decades of neglect by Congress.

The quota limit of ~120,000 per year for legal workers is about 10-100 times too small.  This is where our ~10-11 million illegal immigrant population grew from,  over those same decades of neglect. 

The jobs are here,  the workers that will do them are from down there.  They have to come just to eat,  legal or illegal.  A guest worker visa itself is not a path to citizenship,  but it need not preclude the holder from seeking such. 

Our border agents,  tied up trying to stop so many illegal guest workers,  cannot also deal so effectively with the cross-border smugglers.

A whole-border wall “fixes” little to none of this,  because any wall can be defeated.

Why not just adjust the guest worker visa quotas to realistic levels (and staff up to track them properly,  not done now),  thus freeing a great many border agents to deal with the smugglers? 

Smugglers (of People or Drugs)

Government data clearly shows most drugs come through designated and manned ports of entry,  hidden among legal cargoes.  A whole-border wall does nothing to fix that problem.

Some drugs come by sea or air.  A whole-border wall does nothing to fix that problem.

Asylum seekers do resort to illegal smuggling,  because legal entry has been made so difficult.  And the recent Trump administration rules changes have worsened that entry difficulty,  further incentivizing their resort to smugglers. 

Our border agents are too tied up dealing with asylum seekers and illegal guest workers to deal adequately with cross-border smugglers. 

Why not just fix the two problems (asylum seekers and guest workers) that are sopping up all the agent manpower?  Turn them loose upon the smugglers. 

About the Whole-Border Wall

This was a campaign promise to wall-off the entire southern border.  Sounds great as a sound bite,  but it won’t fix the real problems.

Such a wall is ineffective because defeating a wall is easy:  ladders,  ropes,  gloves,  shovels,  bolt cutters,  and saws-alls are all very much cheaper than fences or walls of any type.

Border walls are ineffective against the majority of the drug smuggling,  because they cross at ports-of-entry,  not all along the border.

According to government statistics,  there are no terrorists at the southern border.  There have been a small handful apprehended at the northern border,  but by far the most were apprehended at airports.  Despite the false claims,  terrorists are no reason for a whole-border wall.

Building a whole-border wall requires the government taking private lands by eminent domain.  This is extremely unpopular in Texas,  especially among border region landowners.  This is true regardless of party affiliation,  according to the polls.  As well it should be. 

Repairs or additions to existing border fencing are fine,  but there is quite obviously no need for a whole-border wall. 

About the Shutdown

The government shutdown is merely a way to hold hostage some Americans,  the US economy,  and US public safety,  in order to fulfill a campaign promise in a highly-visible wayThis does increasing damage the longer it goes on.

The President cannot do this alone.  Key members of Congress must collude with him,  for this damaging grandstand play to be successful.  They do damage to the country for party advantage.

The House and Senate both already had funding bills that contained border security funding, including for barriers.  There are enough votes in the House and Senate,  to pass one of those existing bills,  and end this shutdown,  right now!  There are actually enough votes to override a Presidential veto.

Trump reneged on his promise to sign one of those existing bills,  because of bad publicity he got from the gadflies on Fox-and-Friends and talk-radio.  That is no justification for damaging the country.

In the Senate,  majority leader Mitch McConnell will not allow any of the Senate funding bills to come to a vote without pre-approval from Trump.  Since when does the Senate need approval from the President to do its business? 

This is McConnell prioritizing party advantage above providing for the good of the country.  Is that what any of you really want?  Damaging the country to score political “points”?

What Are the Right Things To Do?

The opposition in Congress cannot give in to hostage blackmail from the White House that is damaging and endangering the country.  Yielding only ensures this evil behavior will be repeated in the future.

The House and Senate need to pass one of the existing funding bills as fast as they can,  and end this travesty.  They should have done this the very first week

If necessary,  the House and Senate should quickly override a Trump veto.   The votes are there to do it.


If Mitch McConnell will not cease obstructing these votes,  then the Senate needs a new majority leader,  one who knows that the good of the country outweighs other considerations!   

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postscript
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For $5 billion,  you could hire more than 20,000 new immigration judges and their support staff,  border agents and their support staff,  and beaucoup other paperwork officers and clerks to track visas.  This assumes on-average about $200,000 each,  annually,  to cover salaries and benefits.  Some cost more,  some less,  but that’s a decent ballpark cost figure for estimates.  Over-20,000 is a whole lot more people than we have now,  working these problems.

That means you could staff up to take care of asylum seekers properly,  while cutting the processing delay to near zero,  and thus reducing the need for proper holding facilities.  It means you could staff up to issue a whole lot more guest worker visas,  and actually track them to ensure proper renewals and no overstays.  And,  the DACA problem goes away within a generation,  once these guest workers are legal. 

Doing those two items correctly frees up a whole lot of border agents to deal with the smugglers a whole lot more effectively!  And very likely with some money left over to upgrade or replace existing border fences,  and to add some more,  where such is actually needed. 

Together,  that solves all the problems,  and without doing an inherently-defeatable whole-border wall,  and taking people’s lands to build it (which takes years to accomplish)!  And so doing this right actually solves the problems quicker,  to boot!

Now,  the facts are quite different from the propaganda,  which is why I wrote this article.  The sane things to do are quite different from the campaign slogans and sound bite crap we are being fed.  

Why on God’s green Earth would anyone with two working brain cells to rub together,  actually believe we need hundreds more miles of tall wall along our southern border?  When we can do far more,  for less money,  and get a better result?

I recommend you apply “grassroots term limits” and vote all these corrupt incumbents out,  who have been damaging our country for nothing but political points scored.  That applies to both parties,  the current senate majority leader,  and the current occupant of the White House.  They all get corrupted by the big money that infects our system everywhere,  within about 1 term in office.  We don’t need that.  No more.

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Update 2-3-19 The end of the shutdown came with a bipartisan passage of yet another short continuing resolution.  Trump signed,  I presume so as not to endure the bad publicity of a veto override.  Yet this is just another in a long series of very short-term continuing resolutions;  this one only 3 weeks!  What looms quickly is another shutdown,  or else serious abuse of the emergency powers law. That is entirely unacceptable!

One of Congress's mandated jobs is funding the government for the fiscal year,  not just temporarily!  This chronic continuing resolution process is nothing but abusing the process to play party politics instead of doing the people's business.  It damages the country.  It is clear evidence of mis-prioritizing party advantage above the public good!

In my opinion,  that is a crime against all Americans. It has gone unpunished for decades.

As I said above,  we voters already have the "grassroots term limits" power.  Vote these people out of office!  It really doesn't matter who,  or what party,  replaces them!  Just exactly how could you do worse than what you have now?

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