Water on Mars?

This is only meant as a funny.  The caption is something like "first discovery of water on Mars".

Bounding Analysis for Lunar Lander Designs

I did this as a "clean sheet" bounding analysis.  Friends I correspond with have asked repeatedly how a lunar base might be established,  and with what.  I know the most about Spacex rockets and capsules,  but actually fleshing out these designs could use anybody's existing equipment,  in whole or in part.  The challenge I now throw out to them is to design something within these limits.

Bounding Analysis for Lunar Lander Designs  

GWJ  3-18-17 completed 3-18-17

The scenario here is a lander delivered “neat” to lunar orbit as an unmanned item.  A crew will arrive separately to rendezvous with it in lunar orbit.  The plane of that orbit is presumed to be very close to the ecliptic.  Orbital direction is retrograde,  in accordance with the figure-eight patched-conic trajectory used during Apollo.  The delta-vee to land one-way is 1.68 km/s.  For design purposes,  a few percent higher is used to provide a little margin:  1.75 km/s. 

The lander is delivered “neat” to lunar orbit,  meaning the rocket that takes it to the moon must do the “burn” to put it into lunar orbit.  The total rocket design delta-vee from the surface of the Earth to do that is at most 12.4 km/s,  when the first 8 km/s to Earth orbit is factored for drag and gravity losses by 1.05.  This is very close to the surface launch for a more-or-less worst-case slow trajectory to Mars,  which is about 12.1 km/s,  factored the same way.  That way,  the tonnage sendable onto a Mars transfer trajectory is almost the same as what can be delivered into lunar orbit,  for our purposes here. 

Descent Design Requirements

Spacex lists on its website that its Falcon-Heavy can send 13.6 metric tons to Mars,  flown fully-expendably,  for about $90 M launch price.  This heavy-lift booster hasn’t yet flown,  but it should fly this year (2017).  Reducing that payload slightly for the slightly-higher delta-vee to lunar orbit,  call that a max payload to lunar orbit of an even 13 metric tons.  

For the descent stage,  ready to fire in lunar orbit,  we are looking at an ignition mass of 13,000 kg maximum,  and a required design delta-vee of 1.75 km/s.  Propellants should be storable,  since days to weeks,  even months,  in space (or on the moon) are contemplated.  With nozzles designed for vacuum,  and assuming NTO-MMH propellants,  a delivered Isp = 335 sec is quite realistic.  Engine thrust/weight ratio of 100 Newtons-of-thrust per Newton of engine Earth weight seems feasible.  

Thrust to ignition Earth weight ratio should just barely exceed lunar gravity’s pull,  so that plenty of thrust margin is available at burnout weight:  0.2 seems “reasonable”.  We’d like the vehicle acceleration at burnout to be less than 1 gee,  preferably under 0.5 gee,  to keep the ride from being too rough,  and to limit throttleability requirements to feasible values.   

The propellant tanks will need a sun-reflective surface and some insulation,  plus electric in-tank heaters,  on a single-hull tank.  That means the descent stage propellant tankage will not be quite as lightweight as that of an expendable booster.  Just considering the tankage alone,  a 95-5 split of propellant to tank masses seems reasonable to assume (Wp/Wt = 95/5 = 19). 

The rest of the stage structure must bear the thrusted flight maneuvering loads carrying as large a payload as possible,  plus incorporate a set of broad-span landing legs,  and some means of unloading large items (ramps,  crane,  etc.).  An inert structural fraction for the stage near 15% should cover all of this.  That fraction does not include tank inerts or engine hardware.  Those get figured separately,  and then added to determine an overall stage inert mass fraction. 

The objective here is to determine max payload mass within that ignition mass limitation.  That payload can be either (1) cargo delivered one-way,  or (2) an ascent vehicle carrying minimum crew and cargo weight.  They mass the same,  though. 

Sizing a “Clean-Sheet” Bound on the Descent Stage

Exhaust velocity is rather accurately estimated as Vex, km/s = 9.8067*(Isp, sec)/1000.  That and the design delta-vee value combine to determine stage mass ratio MR = exp(dV/Vex).  The required propellant fraction (of ignition mass) is Wp/Wig = 1 – 1/MR.  The corresponding fraction for tankage inerts is Wt/Wig = (Wp/Wig)/(Wp/Wt).  The corresponding engine inert fraction of ignition mass is We/Wig = (T/Wig)/(T/We).  The rest of the stage structural and equipment inerts is represented by the 15% figure.  These total together for the overall stage inert fraction. 

Payload fraction of ignition mass is just 1 minus the propellant fraction and minus the sum total inert fraction.  Once you specify an absolute payload mass,  that determines ignition mass,  from which all the component masses are determined by their fractions.  That finalizes the weight statement.  For this bounding exercise based on Falcon-Heavy delivery,  those results are in Figure 1.  

Figure 1 – Limits for Descent Stage,  One-Way,  Falcon-Heavy Delivery to Lunar Orbit

Payload is 5.372 metric tons.  This could be all cargo,  or it could represent a crewed ascent stage.  If cargo,  that’s $90M/5.372 metric tons = $16.8M per metric ton delivered to the surface of the moon.  Actually,  you design to a slightly-smaller payload mass,  because of all the uncertainties.  There is always the unexpected outcome,  when sizing vehicles like this “from scratch”.  The weight margins don’t have to be all that large,  because I already put that into the design delta-vee figures. 

Ascent Design Requirements

The same propellant and tankage choices are presumed.  The same engine T/We is assumed.  A slightly-higher T/Wig = 0.3 is assumed,  to accelerate “smartly upward” against lunar gravity.  Stage inert fractions can be lower since no unload equipment or landing legs are needed.  However,  these inerts are likely higher than a typical booster rocket (5%) because of the protective cabin surrounding the crew,  the docking hatch,  and the instruments and controls they must use.  I simply assumed 10%. 

This ascent stage must ascend to lunar orbit (requiring 1.68 km/s),  and also maneuver to rendezvous with the crew return craft left in lunar orbit.  It therefore needs more design delta-vee than the descent stage.  Call it 2.0 km/s,  for a kitty of 0.3 km/s to cover maneuvering and the unexpected. 

Its maximum ignition mass cannot exceed the descent stage payload capability of 5372 kg.  Prudence dictates very slightly less.  Call it 5360 kg for design-bounding purposes. 

Sizing a Clean-Sheet Ascent Design to Fit the Descent Stage

All the calculations and equations are basically the same as before.  I simply used the same spreadsheet with different numbers.  The results are given in Figure 2.  Deliverable “payload” is 2235 kg,  which would be suited crew plus a few of days of life support,  plus any samples sent back to Earth.

Figure 2 – Limits for Ascent Stage,  One-Way,  To Fit Descent Stage That Falcon-Heavy Can Deliver

For the sake of argument,  use 80 kg per person body weight,  and 120 kg for a surface EVA-capable pressure suit.  That’s 200 kg per person.  Set food,  water,  and breathing oxygen supplies to 100 kg to cover an unexpectedly-long rendezvous interval of several days.  That’s 300 kg allotted per person.  There’s “room” for 7 such masses in the payload. 

If this were 6 crew,  there’s room for around 300 kg of samples or return cargo.  If the crew is 5,  there’s room for about 600 kg of samples or return cargo,  and so forth.  But the point is,  there’s room for a much larger crew than Apollo had.  That’s partly the difference in technologically-achievable storable propellant performance,  and in structural technologies,  since the 1960’s.  The rest is landing without unknown obstacles in your path,  which is what happened on Apollo 11,  nearly depleting its propellant. 

How This Can Be Used

The one-way cargo-only variant can be used at $90M a shot to deliver 5.36 metric tons of cargo to the moon ($16.8 M/delivered metric ton).  Several could be sent to the same site.  Some of these could be the modules from which some sort of surface habitat could be assembled.  The rest could be the supplies,  equipment,  and surface rover vehicles needed to operate that base. 

The manned lander conforms to the same 5.36 metric ton weight limit.  If crew were 3,  then 1200 kg of surface supplies could go down with them.  If crew were 2,  then 1500 kg of cargo could ride down.  Reducing the ascent load just increases the rendezvous maneuver capability upon returning to lunar orbit,  a very beneficial safety factor.    

Say,  we sent 9 of these to the moon:  6 cargo-only landers and 3 landers with manned ascent stages,  each with a crew of 2 and 1500 kg of cargo on board.  That gives us three ascent vehicles on the lunar surface ready to use,  when the entire crew really only needs one to return.  Added safety,  that is. 

That’s a total of 32.16 tons delivered with the cargo landers,  and 4.5 more tons sent down with the manned landers,  for a total crew of 6.  Assume simply for the sake of argument that the surface habitat requires 20 tons.  We need to reserve 0.6 tons of supplies for the crew to ascend.  Assume two rovers,  each 1 ton.  Assume one electric backhoe-like device,  at 2 tons. 

36.66 tons total delivered cargo,  less 20 ton habitat,  4 tons for vehicles,  and 0.6 tons for ascent supplies,  leaves 12.06 tons allocatable for surface stay supplies and other equipment items.  At a nominally-assumed 10 kg life support per person per day for 3 months,  then about half that 12 tons is something other than life support supplies.  Also nominally,  3 months of life support supplies for a crew of 6 is pretty close to a lander’s deliverable payload at 5400 kg.  I tried to overestimate this requirement. 

Looks to me like there is very good potential for establishing a fairly substantial lunar experiment station,  temporarily occupied for a considerable time (at least 3 months).  This requires 9 Falcon-Heavy fully-expendable launches for the landers,  plus one more to send the crew in a crew Dragon (with its trunk modified to carry propellant,  something not addressed here),  for $900 M in launch costs.  If launch costs were 20% of the program that develops these vehicles and the surface equipment,  total program cost to put a small base temporarily on the moon would be in the ballpark of $4.5 B. 

Launching another cargo lander every 3 months or thereabouts brings the supplies to keep that base permanently occupied at crew size 6.  Maybe switch out crews yearly,  by adding a crewed Dragon to lunar orbit along with a fresh manned lander to take them down to the surface.  That’s a total of 6 Falcon-Heavy launches per year to maintain a continuous presence at the base.  That’s $540M per year to maintain the base after it is built,  plus the costs of keeping the necessary vehicles and equipment in production.  Development is complete,  so call launch costs ~50% of continuing program costs. 

About $4.5 B to establish a 3-month-capable,  6-man base on the moon,  and about $1B/year to keep it continuously manned and operating is just not very expensive as space ventures go!  This analysis is based on the use of a commercial heavy lift rocket that is far less expensive to use than NASA’s SLS,  and which will also be far more available for routine use multiple times per year,  than NASA’s SLS ever can. 

Blue Origin is also planning to get into this kind of lunar capability with its New Glenn rocket.  Between them and Spacex,  putting a base on the moon looks to be quite feasible and quite affordable.  This could provide the bootstrap start needed to begin doing something useful,  or for profit,  on the moon. 

Final Remarks

This kind of experiment station allows evaluation of low-gravity effects upon health versus the zero-gravity effects that we are familiar with in Earth orbit.  It allows a place to experiment with increasingly-capable recycling life support systems.  It allows a place to experiment with meteoroid and radiation protection by regolith cover.  It allows a place to experiment with ways and means to overcome contamination and wear issues with very-fine-but-sharp-edged dust particles.  All these are needed to visit Mars or the asteroids,  and are available on the moon “close by” in case of trouble. 

The same base allows experimentation with ways and means to dig and drill deep in a harsh environment.  It allows experimentation with the recovery of mineral resources.  It allows experimentation with how to establish roads under such conditions,  so that future long-distance surface transport becomes feasible.  These things are needed for establishing useful and prosperous industrial applications on the moon and Mars,  and to some extent the asteroids. 

This is the kind of thing we should have attempted to close-out Apollo,  had a useful lunar presence been the goal,  instead of “flags-and-footprints”.  It is still a good rationale for returning and doing something very much like what I described here,  as a first step. 

Addendum:  Crew Dragon Modified to Leave Lunar Orbit

The “design” trajectory to reach lunar orbit is pretty much the same as was used for Apollo decades ago.  A direct launch from Canaveral into low Earth orbit more or less eastward at low inclination (the part requiring factoring ideal delta-vee for gravity and drag losses),  followed by a burn to escape onto the lunar transfer trajectory,  and a final upper-stage burn to place the payload into a retrograde orbit about the moon.  The worst-case total rocket design delta-vee for this is just about 12.4 km/s (factored),  and worst-case 0.8 km/s to leave lunar orbit onto a trajectory home.  See Figure 3. 

If we stay under 13 thrown metric tons,  the Falcon-Heavy should have enough delta-vee capability to put that 13 tons into lunar orbit,  same as the lander designs just bounded above.  The problem is then leaving lunar orbit with enough propellant reserve to cover attitude control and a powered landing on land back on Earth (Spacex’s preferred mode).  Attitude control consumption should be modest,  but we might need around 0.5 km/s capability to land safely on Earth,  where capsule terminal fall velocities are only around half a Mach number.  0.8 + 0.5 + small change is close to 1.35-1.4 km/s delta vee capability demanded of the Super Dracos on crew Dragon.  It simply does not have that much capability without extra propellant added in the trunk,  and connected to the system in the capsule.

Figure 3 – Design Trajectory and Delta-Vee Requirements

Design Requirements for Modified Crew Dragon

Total delta-vee capability 1.35 km/s min,  1.40 preferred.  Maximum spacecraft mass at launch 13.0 metric tons.  Minimum crew 3.  I have a spreadsheet model already constructed for this purpose,  which I proceeded to run again for these exact numbers.  Masses for the dry weights of capsule and trunk (before modification) are my best guesses,  but their sum matches published data. 

The modification is to install more tanks of NTO-MMH propellants in the trunk,  to a maximum of the 3000 kg quoted cargo capacity for that trunk.  I estimated propellant-tank mass split as 95-5 or a 19:1 ratio,  same as for the landers.  I did not estimate volumes,  although there are 14 cubic meters available in the trunk for this. 

Results That Bound the Design

These are shown in Figure 4.  Payload mass is limited more by the 13.00 ton thrown weight than the 1.35-1.4 km/s delta-vee requirement.  That payload mass is 1760 kg. 

The per person allotment we used for the lander was 200 kg person-plus-suit,  and 100 kg of packed life support supplies.  The life support supplies are probably a bit of an overkill,  so 1760 kg ~ 1800 kg,  and 1800 kg / 300 kg/person is crew = 6 max.  Slightly less actually.  Call it no more than 5 crew at a time,  plus life support supplies,  and no more than about 150 kg of equipment or cargo in the capsule with them,  for the trip to the moon. 

Having the extra delta-vee means we can carry 6 crew,  even 7,   home.  That is a good safety bonus.  Crew Dragon is supposedly rated for the same cargo home as cargo Dragon (3000 kg),  so we are well within that limit. 

This was accomplished by adding 2800 kg propellants to the trunk,  which also adds 147 kg of tank inerts to the trunk inert weight.  That leaves a smidge for any extra plumbing before we hit the 3000 kg limit. 

The only remaining question is for how long a crew Dragon can be parked in lunar orbit before the crew that needs it must come back to it and fly it home.  There are limits to lifetime allowable parked in space.  Perhaps this can be made into some number of months to a year,  given some experience flying the capsule design in Earth orbit.  If a year,  then the lander plan given above is quite feasible just as it is laid out.  If not,  we’ll have to switch out crews on the moon at 6 months,  perhaps.

Figure 4 – Modifying Crew Dragon Into Lunar Orbit Dragon for Falcon-Heavy Launch

Final Remarks

With these two bounding analyses,  I have shown how it is possible to ship 13-ton lunar cargo and crew landers to the lunar orbit with Falcon-Heavy as the launch rocket.  I have also shown how it is possible to ship crews to lunar orbit with the same rocket and a 13 ton modified crew Dragon that has 2.8 extra tons of propellant in its trunk,  connected to the Super Draco thruster systems in the capsule. 

The cargo landers deliver slightly over 5.3 tons to the surface.  The crew landers have a 5.3 ton ascent stage that could carry as many as 6 crew back to lunar orbit. 

At only $17M/delivered ton,  building a practical small experiment station that is permanently occupied becomes easily possible,  at a price far below what was experienced doing the Apollo “flag-and-footprints” stuff during the cold war. 

What makes this feasible is a heavy lift rocket of adequate size to put 13-ton payloads into lunar orbit,  and at a commercial launcher’s far lower price.  This is true flying the rockets fully-expendably.  This capability should become available within the next 1-2 years. 

All that is needed from a vehicle development standpoint is the two versions of the lander designed to these bounding limits,  and then developed and made ready for use.  These share a common descent stage.  That should help lower costs and development time. 

Adding propellant capacity to crewed Dragon with tankage in the trunk is not so much development work,  more of a routine modification that can be tested all-up in Earth orbit,  to make it ready to use.
 

We’ll need a 2 or 3 seat open electric rover car that weighs no more than a ton.  Between the Apollo rovers and the recent Mars robot rovers,  this should not be a major development item.  

Development,  yes,  just not a “biggie”.  Same for a 2-ton electric front-end loader. 

The hardest nut to crack is a surface habitat that can be assembled from modules that fit within the 5 ton lander payload capacity,  and that can be erected by men on foot in spacesuits with hand tools.  The idea is to assemble it in an excavation done with the front end loader,  and then bury it at least partially with that front end loader. 

This is the kind of thing that could be done within 1 or 2 presidential terms,  which would net returns orders of magnitude greater than Apollo,  for costs orders of magnitude less than Apollo.  

Reverse-Engineered "Dragon" Data

Reverse-Engineering What the Versions of “Dragon” Can Do       

GWJ       2-17-17                 updated 3-5-17

Sources:  Spacex’s website and the Wikipedia articles on cargo Dragon,  crewed Dragon,  and Red Dragon.  There is also DragonLab,  which is a very close variant of cargo Dragon.  These give dry weights for the spacecraft that seem to include the associated trunks,  except in the case of Red Dragon,  which is listed in a very sparse article as “6.5 ton plus payload up to 1 ton”.   Comments made in public by Spacex have indicated the possibility of more than 2 metric tons payload to Mars for some time now. 

Cargo Dragon:  The Wikipedia article lists dry mass as 4200 kg,  and speaks of a chute drop test at 5400 kg that includes a max cargo weight of 2500 kg.  Propellant quantity for the Draco thrusters is no longer on Spacex’s site,  but was once listed as just about 1290 kg.  The capsule has a jettisoned nose cone fairing for ascent,  for which a wild guess is 50 kg. 

The ocean landing test configuration would be capsule dry mass plus max rated cargo,  plus some propellant residual if not jettisoned after entry and chute deployment.  Being toxic,  they should be jettisoned before recovery is attempted by humans.  I assumed zero propellant residuals,  so that the actual capsule and trunk dry masses could be determined in this way:

Both the website and the Wikipedia article list max cargo “up” as 6000 kg,  with at most 3000 kg in the capsule,  and with 3000 kg unpressurized in the trunk.  Max “down” cargo is listed as 2500 kg in the capsule,  with up to 3000 kg of waste in the trunk to be destroyed on entry.  Max cargo available to be carried to the ISS is listed in the Wikipedia article as 3310 kg,  presumably a max of 3000 kg in the capsule,  and the rest in the trunk.   The station’s arm is used to unload items in the trunk.

Weight statements for cargo Dragon can now be estimated to the accuracy that trunk dry mass estimate is accurate,  and that the nose cone mass can be guessed.  For three possible cargo loadouts these are:

Compare the launch weights above with Falcon-9 capability to LEO from Spacex’s website.  If flown as an all-expendable launcher,  the rocket can send 22.8 metric tons to LEO,  and only as a guess probably something close to perhaps 15-17 metric tons to ISS.  All the cargo Dragon estimates shown above fall well within that capability,  at none over 11.5 metric tons.  Whether the booster core is recoverable at 11.5 tons is just not determinable (the website does not list those reduced payload limitations).

One of the things in the weight statement is the set of ignition and burnout weights for the capsule-only,  no trunk.  Cargo Dragon is not operated that way,  however!  It retains the trunk until after the reentry burn.  So it is capsule-plus-trunk ignition and burnout weights that we are really interested in. 

To get those ignition and burnout weights,  you add the capsule-only ignition weight and the total loaded trunk weight for capsule-plus-trunk ignition weight,  from which you delete the propellant for burnout weight.  This leaves out the nose cap,  which was already jettisoned during ascent.  

The cargo Dragon has 18 Draco thrusters arranged in two pods of 4 and two pods of 5 within the outer mold line.  These provide attitude control and maneuvering delta-vee,  plus re-entry delta-vee.  Each Draco is about 90 lb thrust (400 N).  These burn NTO-MMH,  for which one might assume Isp = 335 sec for a “good vacuum” engine design,  meaning a long bell for high expansion ratio.  The corresponding mass ratios (MR) and max theoretical delta-vee capabilities for capsule-plus-trunk are:

Crewed Dragon (Dragon v2):  This is the same basic capsule pressure shell and mold line,  modified for four protruding pods,  each pod containing two Super Draco thruster engines and four Draco thrusters for attitude control and minor maneuver.  The Super Dracos are listed on the Wikipedia article as 16,000 lb thrust (71 KN) each.  They use the same propellants as the Dracos.  Spacex’s website lists the eight total Super Dracos as having 200,000 lb axially-directed thrust (890 KN).  Older versions of the site listed the propellant load as just about 1890 kg.

The Wikipedia article lists dry weight as 6400 kg,  which apparently includes an empty trunk.  This capsule has the same chutes,  a retained reusable nose cap,  crew life support,  crewed interior seats and fitments,  and landing legs.  The trunk is of similar size,  but arranged with conformal surface solar panels instead of folding solar panel wings.  It does have four aerodynamic fin surfaces for aerodynamic stability during crew escape situations.  There is no information available anywhere I can find by which to separate the trunk dry mass from the capsule-plus-trunk dry mass. 

Crewed Dragon operates in space as capsule-plus-trunk,  until after the reentry burn,  when the trunk is jettisoned.  The Wikipedia article lists exactly the same cargo masses and volumes as for cargo dragon.  Unlike cargo Dragon,  crewed Dragon uses the chutes only as a safety backup landing method,  or for landing in the ocean.  Its intended mode is a propulsive landing on dry land with the Super Draco engines,  no chutes at all.  Because of this,  both the capsule-only and capsule-with-trunk max theoretical delta-vees are of interest.  Note however that you cannot achieve both simultaneously,  because there is only one propellant supply to be used for both purposes! 

The best I could do was to simply assume the two trunks were comparable mass in spite of the design differences.  The uncertainty in the resulting data is dominated by that assumption.  Again,  I assumed Isp = 335 sec for an exhaust velocity of 3.285 km/s.  Using cargo Dragon’s trunk mass, the crewed dragon capsule dry mass (which includes the reusable nosecone) is:

This capsule-only dry mass is slightly more than 2 tons higher than cargo Dragon,  but there are the eight Super Draco engines,  an uprated heat shield,  landing legs,  life support,  and crew seats and fitments to consider,  so it is “reasonable”.  From this, one can estimate the same sort of weight statement breakout already reported for cargo Dragon,  including both capsule-only and capsule-plus-trunk ignition and burnout weights.  I did this for only one crew/cargo value,  chosen to approximate a capsule-only delta-vee of 0.7 km/s to compare with Red Dragon.

These figures show comparable values of capsule-plus-trunk delta-vee to cargo Dragon’s ~0.5 km/s,  which is realistic,  considering crewed Dragon is a derivative design,  operating in the same capsule-plus-trunk configuration.  The slightly-higher capsule-only figure is to compare with crewed Dragon’s unmanned derivative Red Dragon (for one-way probes to Mars).  Note that this would vary significantly as crew/cargo is adjusted.  Under the assumptions of 100 kg person,  100 kg suit,  50 kg air and water,  we are talking about 7 crew plus 1050 kg cargo in this 2800 kg loadout.  The weight to launch falls well within the Falcon-9’s LEO capability,  being just about like the heaviest cargo Dragon presented above.

Red Dragon:  This is the crewed Dragon with the crew seats and fitments,  life support,  and chutes removed,  and some equipment racks installed.  The heat shield is reduced in thickness as well.  Since the vehicle is not to be reused,  a jettisonable nose cap like that of cargo Dragon is assumed.  It will need some sort of trunk for launch and for electricity during the trip,  but this is jettisoned before Mars entry.  Course correction burn is assumed trivial,  so that essentially the entire propellant load is available for powered landing on Mars. 

There are no available data for the masses of any of the change items just discussed.  All that is available are wild guesses and educated guesses.  The lighter heat shield I estimated as a reduction from 8 cm thick to 6 cm thick,  on a flat circle 3.7 m diameter,  and a specific gravity of ~0.3 for PICA-X.  I just rounded off to the nearest 10 kg.  It’s just too rough not to round off like that.    

I have just assumed the same trunk mass as I used for cargo and crewed Dragons.  Trunk mass dominates the uncertainty,  being the largest item.  I simply took the crewed Dragon estimated dry mass,  and subtracted things.  Those guesses are listed in this estimate for Red Dragon dry mass:

If I load this vehicle with 1 ton of cargo and 1890 kg propellant,  then mass at entry is 7640 kg,  which is not far at all from the 7500 kg indicated the Wikipedia article!  1 ton of cargo is what is indicated as deliverable to Mars in that same article.  (Some public announcements indicate that 2-4 tons are actually under consideration at Spacex.)

Using these figures,  the weight statement for Red Dragon can be roughly estimated.  What is of interest here is the capsule-only delta-vee,  as a function of cargo delivered to the surface of Mars.  Bear in mind that an utter minimum delta-vee capability for powered landing will be near 0.7 km/s,  the Mach 3 point coming out of atmospheric entry hypersonics.  There’s very little in the way of gravity and drag losses to correct the theoretical delta-vee in this scenario.  The error is less than the uncertainty in the basic requirement.

The 0.7 km/s figure is pretty rough,  that being 3 times the nominal speed of sound in the Martian atmosphere at something like 5 km altitudes.  This could vary quite a bit.  In order to successfully land reliably,  you actually need a little more delta-vee to cover final maneuvering around obstacles. 

The Mach 3 point is also a bit arbitrary,  that being only the definition of min-hypersonic for blunt objects.  Prior probes deployed chutes at local Mach 2 to 2.5,  although they did this much higher up (15-25 km altitudes or more).  Waiting to lower speeds lets you penetrate to lower altitudes,  while heavier items also penetrate to lower altitudes,  simply because of higher ballistic coefficients. 

The 0.7 km/s “requirement” I use here is thus just a figure of merit,  although it is actually in the ballpark of the true requirement.  

These numbers are too rough to judge “for sure”, but it looks like 1 or 2 metric tons should be easily deliverable to Mars with Red Dragon,  just like Spacex has indicated.  These numbers say 3.2 tons is getting to be quite marginal,  but that would actually depend upon what the true landing delta-vee requirement really is.  Note that the requirement would vary with location and season across Mars,  as that atmosphere is much more variable in its density than is Earth’s. 

Spacex’s website lists Falcon-Heavy as able to send 13.6 metric tons to Mars,  flown fully-expendably.  All these configurations fall within that capability.  Even at 4 tons cargo,  the weight to launch would be just about 12 tons,  which still falls within the launch capability.  Therefore,  it will be landing delta-vee that sets the payload deliverable to Mars!  That may explain the “extra propellant” remark found in the Wikipedia article.  Whether 11 tons at launch is small enough to recover the booster cores is unknown. 

DragonLab:  this is cargo Dragon with an instrument bay between the pressure shell and the outer mold line.  The door on this bay opens to space,  and recloses before entry.  It is otherwise the same as cargo Dragon,  so no separate analysis is done here.    Use my figures for cargo Dragon to represent DragonLab.

Using variants of Red Dragon as unmanned one-way probes elsewhere:  This depends on the delta-vee requirement to land,  relative to the vehicle capability.  For airless destinations with direct landings from the interplanetary trajectory,  several-to-many percent above the body’s escape speed is a figure of merit.  Example:  Mars escape 5 km/s,  typical direct entry speed 6 to 7 km/s.  Call it 1.5 V_esc as a typical figure of merit.  Remember that this is only a very crude estimate,  unlikely to be correct!

Potential destinations include the moons of Jupiter and Saturn,  the asteroids,  and our own moon.  Mars and Venus have atmospheres that allow aerobraking to a propulsive landing,  as does Titan at Saturn.  Mercury requires all-propulsive landing,  as does the moon.  Values of escape speed and surface gravity strength follow.  I did not include Venus because landed vehicle lifetime would be too short,  if it made it down to the surface at all.  I also did not include Earth itself.

The first thing apparent from the list is that any of the asteroids,  Titan,  and Mars all seem to be within reach of Red Dragon on a Falcon-Heavy,  just as it is.  The significant atmospheres of Mars and Titan make aerobraking feasible,  the rest are airless or so tenuous as to make aerobraking infeasible.  Like the Earth’s moon,  the moons of Jupiter seem out-of-reach,  due to escape speeds that are too high.

As a one-way probe destination,  Earth’s moon is interesting on its own.  Key here is getting into lunar orbit using the upper stage of the launch rocket,  without using any of the Dragon’s propellant.  As it turns out,  that delta-vee requirement for the launch rocket (no more than 12.4 km/s) is very similar to that for sending things to Mars (at least 12.1 km/s).   Those figures include 5% gravity/drag losses on the first 8 km/s of that delta-vee.   From there it takes 1.68 km/s to make a powered landing.  That’s out-of-reach for Red Dragon without considerable extra propellant. 

It might be more desirable to instead enter lunar orbit with a crewed Dragon,  and let them rendezvous in lunar orbit with a separately-sent lunar lander.    

The lander must descend and ascend to a delta-vee of essentially the orbit velocity each way, for an utter minimum of something like 3.36 km/s plus a tad for gravity losses.  At the moon,  the crew needs a minimal place to ride,  not a full capsule,  but the stage does need landing legs.  It needs to carry surface stay gear and a rover,  as well.  That design is not explored here.

As for the missions to the outer moons,  there needs to be a fairly-large propulsion stage added to the Red Dragon.  It seems like the Dragon probe assembly could be sent to Earth orbit on a Falcon-9,  and the propulsion stage sent there with a Falcon-Heavy,  to be docked together in orbit,  and launched on its mission from there.  Very much better information on velocity requirements is needed to size such an exploration stage design.  That is not addressed here. 

Conclusions:  Red Dragon as presently envisioned works for Mars,  Titan,  and any of the asteroids.  The other outer planet moons require a fairly large powered stage added to the Red Dragon to achieve the necessary delta-vee for capture and landing.  The combined weight exceeds Falcon-Heavy capabilities for direct interplanetary trajectories,  so that something other than direct launch to interplanetary travel is required.  Falcon-Heavy is able to fling 13.6 metric tons to Mars,  perhaps 12-13 tons into lunar orbit.

For Apollo-like lunar missions,  crewed Dragon with extra propellant in the trunk (yielding near 1.6 km/s capability) fits one Falcon-Heavy,  and a lander not based on a Dragon fits another Falcon-Heavy.  These need to weigh under 12-13 metric tons to be successfully launched direct from Earth’s surface,  and must rendezvous in lunar orbit.   Red Dragon itself,  as it currently is envisioned,  seems unattractive for one-way missions to the moon,  with a direct landing delta-vee near 2.4 km/s.  However,  whatever added propulsive stage works for the outer planet moons would work at Earth’s moon. 

Addendum 3-4-17:

I looked very crudely at how much propellant to carry in the trunk to enable a crew Dragon to depart from lunar orbit and still have propulsive landing capability at Earth.  This would be for a crew of only two,  with their suits,  and about 500 kg of supplies and samples.  I did actually add propellant tank inert mass to the trunk (about 147 kg),  which requires iterations.  One needs the trunk for electrical power during the trip home to Earth,  so it is capsule-plus-trunk mass ratio and delta-vee that are pertinent. 

Results are given in the Figure just below.  It is slightly over 12 metric tons as launched.  Falcon-Heavy just might be able to deliver this to lunar orbit with a delta-vee of no more than 12.4 km/s,  because Spacex’s website says it can send 13.6 metric tons to Mars (with an estimated delta vee no less than 12.1 km/s).  Those delta vee estimates are my calculations made using Hohmann min-energy transfer ellipses at orbital semi-major axes that produce the largest delta-vee requirements,  and for which I applied 5% gravity and drag loss to the first 8 km/s getting off the surface of the Earth. 

The capsule-plus-trunk needs about 0.8 km/s sec to depart lunar orbit onto a free-fall homeward trajectory to an aerobraking entry.  It perhaps needs about 0.7 km/s capability to cover a propulsive landing on land,  per the basic design of the crewed Dragon.  The total is 1.5 km/s,  for which a few percent margin gets you quickly to 1.6 km/s capability.  The configuration shown in the figure has such capability by adding some 2800 kg pf propellants to the trunk,  in tankage weighing about 147 kg.  That total is 2947 kg,  within the 3000 kg cargo rating of the trunk.  It is the capsule that is lightly loaded at only 2 suited persons plus modest supplies.  Adding a third crew person pushes the total closer to 13 metric tons,  just about the limit that Falcon Heavy could possibly deliver to lunar orbit.  

Addendum 3-5-17:

The same sort of added propellant in the trunk could also be done with Red Dragon.  Note that I have already pretty much defined the maximum propellant already at about 2800 kg.  Red Dragon being similar to crew Dragon,  the capsule-plus-trunk delta-vees will fall far short of what is needed to capture and land on the outer planet moons (1.6 km/s or thereabouts versus 2.5+ km/s).  The capsule-only value indicated in the figure is wrong,  because most of that propellant must be in the trunk.  

I’m not at all convinced that a capsule is the best vehicle by which to send probes to the airless bodies,  including the asteroids,  which Red Dragon could certainly reach.  But Red Dragon should serve very well for probes to Mars,  and maybe Titan.  With aerobraking entry,  a capsule and heat shield are necessary. 

Crewed Dragon could easily become part of a system to send men back to the moon with very few changes from what will start flying this year.  Flyby missions need no changes,  and orbit missions require some extra propellant in the trunk for the Super Dracos.  Landings will require a separate lander sent ahead to lunar orbit unmanned,  for the crew to rendezvous with and utilize.  If such a lander totals under 12-13 metric tons,  Falcon-Heavy could fling it to lunar orbit launched directly from Earth’s surface. 

Final Comments as of Posting 3-6-17:

Enjoy!  These figures may not be exactly right,  but they are pretty close.  I suggest you create spreadsheets to calculate delta-vee in capsule-plus-trunk configurations for cargo Dragon and for crewed Dragon,  and for both capsule-plus-trunk and capsule-only Red Dragon.

Then,  given real delta-vee data to reach a destination,  you can compute for yourself whether the corresponding Dragon configuration can reach it.  Be sure to factor-up astronomical values for gravity and drag losses where those apply.  This is required before you size mass ratios and weight statements.

There's no information out there about it,  but I would hazard the guess that Spacex is already looking at versions of crew Dragon and Red Dragon that have extra propellants in the trunk.  That just makes too much sense for them not to be doing that.

GW