Possibilities for a single launch architecture of the Artemis missions, Page 4: lightweight landers from NRHO to the lunar surface.

 Copyright 2024 Robert Clark

 Congress is becoming increasingly concerned that with the continuing delays of the Artemis missions that China may beat the U.S. back to the Moon:

US must beat China back to the moon, Congress tells NASA.
By Mike Wall 
'It's no secret that China has a goal to surpass the United States by 2045 as global leaders in space. We can't allow this to happen.'
https://www.space.com/us-win-moon-race-china-congress-artemis-hearing

 I had previously proposed correcting an error in the design of Orion's service module that instead of making it larger than Apollo's service module because of Orion's twice larger size, instead made it 1/3rd smaller:

Possibilities for a single launch architecture of the Artemis missions, Page 2: using the Boeing Exploration Upper Stage.

https://exoscientist.blogspot.com/2023/08/possibilities-for-single-launch.html

 The proposal was to give an additional approx. 10 tons propellant to the service module. This would allow the Orion capsule/service module stack plus an Apollo-size lander to be carried all the way to low lunar orbit, not just to NRHO(near-rectilinear halo orbit). 

 This though because of the higher payload may require use of the higher thrust J-2X engine on the Boeing EUS(upper stage) rather than the 4 RL-10 engines now planned on the SLS Block 1B. It's higher thrust would result in a greater payload to LEO and TLI, perhaps to ca. 120 tons to LEO rather than the 105 tons planned to LEO.

This approach requires additional propellant tanks be added to the service module and a change in the EUS upper stage engine to the J-2X. As I discussed in that blog post, it may also require an additional Centaur V sized third stage be added atop the Boeing EUS. This is dependent on what is the TLI(trans lunar injection) payload for the Boeing EUS using the J-2X engine. It may be it can perform the needed TLI payload without an additional Centaur V 3rd stage.

 In any case, I'll propose here an alternative approach to a single launch Artemis architecture without increasing the service module propellant load. This again will use a light-weight Apollo-sized lander with all the components of Orion capsule/Service Module/lunar lander all carried on that one single SLS launch. Because of the lower propellant load on the service module though I'll also send it to NRHO instead of to low lunar orbit.

 Note the NRHO was chosen by NASA as the orbital location because it has a lower delta-v requirement to get there than going to low lunar orbit. Here’s the the delta-v requirements:

 The second group of delta-v’s shows the delta-v to NRHO as 0.45 km/s and the delta-v to and from the lunar surface from NRHO as 2.75 km/s, or 5.5 km/s round trip.

 I’ve seen various numbers for the Orion and service module dry mass and propellant mass. I’ll use 16.5 total dry mass for the Orion+service module together, and 9 tons of service module propellant mass, but only 8.6 tons of this as usable propellant because of residuals.

 Then we'll use 6 tons of Service module propellant to get the Orion/Service Module/lunar lander to NRHO after being placed on TLI trajectory by the EUS, for the 16.5 ton Orion/Service Module dry mass, and 15 tons gross mass Apollo-sized lander with 2.6 tons left over for the return trip.

 We'll need every bit of performance to accomplish the mission within these constraints. So we'll assume we can get a 324 s Isp out of the storable propellant engines on the service module. This is higher than specified for the Orion service modules engines but is doable because of the storable propellant Aestus engine on the Ariane 5 EPS storable propellant upper stage which gets this vacuum Isp. We'll assume we can get this increased Isp by using a larger expansion ratio nozzle or even by swapping out the engine on the service module to use the Aestus engine. Then we get:

324*9.81Ln(1 + 6/(16.5 + 15 + 2.6 + 0.4)) = 510 m/s, or 0.51 km/s, sufficient for placing the stack in the NRHO orbit, where the 0.4 in the equation is for the unburnt residuals.

 Then with the 2.6 tons usable propellant left over for the return trip, after the lander is jettisoned, we get:

324*9.81Ln(1 + 2.6/(16.5 + 0.4)) = 450 m/s, 0.45 km/s, sufficient for the Orion return.

 To increase performance even more we may want to switch even to the RS-72 engine. This is a turbopump-fed storable propellant engine with a vacuum Isp of 340s. It achieves this by using a higher chamber pressure of 60 bar and higher nozzle expansion ratio of 300 to 1 than the Aestus engine. A turbopump engine also has lower residuals, typically less than 1%. A disadvantage is that pressure-fed engines are simpler with fewer moving parts, and so higher reliability, important for an engine to place the spacecraft in orbit and for leaving orbit.

 Now for the ca. 15 ton gross mass lander, because of the higher delta- v needed from NRHO we’ll use hydrolox rather than storable propellant stage. The Ariane 4 H10 hydrolox upper stage had a 11.8 ton propellant mass and 1.2 ton dry mass. We’ll use a 2 ton dry mass of the crew module:

ORBITAL PROPOSES FUTURE DEEP SPACE APPLICATIONS FOR CYGNUS.
SPACEFLIGHT INSIDER
MAY 1ST, 2014
Orbital’s proposal, outlined in this PDF, involves docking a Cygnus spacecraft with Orion to serve as a habitation and logistics module on longer flights. For these missions, the re-purposed Cygnus would be called the Exploration Augmentation Module (EAM). With its current life support systems used to transport pressurized cargo and experiments to the ISS, Cygnus is stated as being already suitable for the long term support of a crew. While berthed to Orion, Cygnus could support a crew of four for up to 60 days. Cygnus also has the capability of storing food, water, oxygen, and waste and features its own power and propulsion systems. The EAM would utilize the enhanced configuration Cygnus, which will begin flying larger cargoes to the ISS beginning with CRS-4 in 2015. An even larger version is also being proposed, featuring a 4-segment pressurized cargo module.

https://www.spaceflightinsider.com/missions/commercial/orbital-proposes-future-deep-space-applications-cygnus/

 Note though the phrasing here is ambiguous. The Cygnus capsule as used as a cargo transport to the ISS contains air, as it would have to for the astronauts at the ISS opening it to retrieve the cargo, but not life support systems. I'm inclined to believe for the usage cited in this article it would be taking life support from the Orion capsule. Then the calculations need to be made for how much mass it would take for life support, thermal management, consumables for an independent crew module.

 Now for the delta-v calculation for our hydrolox lander, we'll assume we can match the max 465 s Isp of the RL-10 engine by giving the Ariane 4 upper stage engine a nozzle extension as used on the RL-10, then we get:

465*9.81Ln(1 + 11.8/(1.2 + 2)) = 7,000 m/s, 7 km/s. This is quite a bit higher than the 5.5 km/s needed for the round trip from NRHO to the lunar surface and back again. But it uses hydrolox propellant so needs extra mass for low-boiloff tech. 

 Low boiloff-tech and long duration hydrolox stages are an important enabling technology. ULA engineers and ULA CEO Tory Bruno have written about this extensively in regards to for example the proposed ACES derivative of the Centaur upper stage. Because of the prior research on low-boiloff tech, an operational version to be fielded in a short time frame to be used on the Artemis missions likely can be done. 

 This shows a single launch mission is doable if going to NRHO, but it is not my preferred plan. A complete orbit around the Moon at NRHO altitude takes about a week, and for the Orion capsule being at NRHO and not low lunar orbit, the lander's crew would have to remain on the Moon about a week before they could return to the Orion in the NRHO orbit. The landers crew module would have to be larger with heavier life support and consumables in this scenario.

 If instead the Orion was at low lunar orbit it takes two hours to complete an orbit and the lunar lander could launch every two hours to rendezvous with the Orion.

 Since the Orion's service module being given an insufficient propellant load is such an obvious design mistake, the preferred route to take would be to correct that error, thereby allowing the missions to take place from low lunar orbit instead of from NRHO.

  Robert Clark

Towards advancing the SpaceX Starship to operational flight: SpaceX should lower the Raptor chamber pressure and thrust level.

 Copyright 2023 Robert Clark

 In the blog post, "Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?", I suggested that SpaceX was able to get the SuperHeavy booster to complete its portion of the ascent in the last test flight by throttling down the thrust on the Raptor engines to less than 75%, and that the Starship upper stage was not able to because its Raptors were run at ~90%.

 Reducing the throttle level reduces the chamber pressure also, from 300 bar to ~225 bar, allowing the Raptor to fire without leaks.

 But if SpaceX lowered the thrust level on the booster to prevent engine failures, why did they not also do this on the Starship upper stage? From the propellant level indicators SpaceX provided on the launch video page, we can estimate the remaining propellant on the Starship just prior to FTS as under 100 tons out of a max 1,200 tons. 

      (The ship is not visible in this image because of its extreme distance just before it exploded.)

 Then we can estimate if the engines were run at less than 75%, the remaining propellant would be in the range of 300 tons. This likely would have been too much SpaceX to reach its goal of getting the Starship to just under orbital velocity:

SpaceX Starship megarocket launches on 2nd-ever test flight, explodes in 'rapid unscheduled disassembly' (video) News.
By Josh Dinner published November 18, 2023
The spacecraft was never expected to reach full orbit around Earth, instead flying on a suborbital trajectory to splash down in the Pacific Ocean off the coast of Hawaii. "We're not targeting orbit today; we're targeting almost orbit," said Siva Bharadvaj, a SpaceX operations engineer, adding that the goal was to "get to a thrust profile similar to what we would need for orbit, but also energy level that the ship would need to dissipate for reentry."
https://www.space.com/spacex-starship-second-test-flight-launch-explodes

  So what would be the payload possible if we ran both stages at ~75% thrust? A rough estimate would be at approx. 100 tons as a fully reusable launcher, instead of the 150 tons now. A problem with that is the estimated number of refuelings for the Starship HLS used as an lunar lander was perhaps 16. But if the max payload was only 100 tons, then the number of refuelings would rise to 24.

 The idea of using so many refuelings for a lunar landing mission has been controversial. Then we'll look at other approaches. Robert Zubrin has noted it can be done in a single launch by giving the SuperHeavy/Starship a small 3rd stage, a mini Starship. 

 Elon Musk once suggested an expendable version of the Starship could have a 30 to 1 mass ratio:

__________________________________________________________________

Elon Musk @elonmusk

Mar 29, 2019

Replying to

@Erdayastronaut and @DiscoverMag

Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible.
__________________________________________________________________

 We'll explore this capability. We'll first look at a case where both the SuperHeavy and Starship are expendable stages. A stage generally gets better mass ratio scaled up, so to begin with we'll start off with an estimated mass ratio for the SuperHeavy as about 30 to 1 with the same propellant load as it is now of 3,400 tons but a dry mass of only 115 tons. 

 But that 30 to 1 mass ratio of the Starship is with just 3 engines since an upper stage does not need enough thrust to lift off from the ground. The SuperHeavy scaled up then would only have 9 engines, but this would not be enough for liftoff. We'll take the number of engines on this version of the SuperHeavy to be again 33. Then we need an additional 24 engines. 

 The thrust/weight ratio of the Raptor is about 140 to 1. Assuming we're reducing the operational thrust level of the Raptor to 75%, its thrust would be 0.75*230 tons-force = 172.5 tons-force. At a 140 to 1 T/W ratio its mass would be 1.2 tons. So the additional 24 engines needed would add 24*1.2 = 28.8 tons to the dry mass: 115 + 28.8 = 143.8 tons. Round it up to 150 tons with the additional plumbing and larger thrust structure required.

 However, with the lower sea level thrust of our 75%-throttle Raptors, the 2-stage SH/SS would have low T/W ratio. We'll also look at a case of adding small 3rd stage of size ~300 tons. Adding a ~300 ton 3rd stage reduces the T/W to barely above 1. Then it would be advisable to add 2 more Raptors to the booster, bringing it to 35. There should be sufficient room beneath the booster for an additional 2 Raptors. 

 The total thrust of the SuperHeavy at sea level is now 35*172.5 = 5,950 tons-force, 58,400 kiloNewtons(kN). We'll use the SilverbirdAstronautics.com payload estimator. This takes the vacuum Isp and vacuum thrust as inputs even for first stage engines. For the lower chamber pressure, ~ 225 bar corresponding to the reduced thrust, we would also have lower vacuum Isp. Call it about that of the Raptor 1 so ~ 350s. Then the vacuum thrust for this version of the Raptor would be (350/327)*172.5 = 184.6 tons-force. For the 35 SuperHeavy engines that's 35*184.6 = 6,460 tons-force, 63,400 kN vacuum thrust. For the dry mass for the 35 engine stage we'll raise it to 152.4 tons.

 For the Starship upper stage, the Raptor Vacuum will still have a vacuum Isp of 380s. However, we are keeping 3 sea level Raptors and 3 vacuum Raptors. Then the average Isp for the upper stage will be 370s. The vacuum thrust calculates out to be 11,500 kN. 

 With the 3 additional Raptor engines being vacuum engines the added weight will be higher than that of just 3 additional sea level engines. We'll take the dry mass  of the new upper stage to be 45,000 kg.

 Then the input page to the SilverbirdAstronautics.com page looks like:

 And the payload results are:

  The expendable payload value of 264 tons is more than the expendable payload of the current SH/SS of 250 tons even-though the new version would be at 75% thrust level. This surprising result must be due the greatly reduced dry mass of both stages. 

 But even more surprising is the payload possible to translunar injection(TLI). In the SilverbirdAstronautics.com calculator this option is indicated by selecting "Escape Trajectory" for "Destination". In the "Hyperbolic, C3" field enter, -1.0. This number indicates how far beyond escape velocity the flight needs. In this case, it's negative since you don't quite have to get to escape velocity to reach the Moon. 

 Then the estimated payload is:

 The payload would be ~75 tons sent to rendevous with the Moon.This is far beyond what Apollo at ~43 tons or SLS even in its later Block 2 version at ~46 tons to TLI could do as a single launch architecture. This would allow wide latitude in how you would design the in-space stages and lander to reach the surface of the  Moon.

 A Three-Stage SuperHeavy/Starship/mini-Starship.

 A basic principle of spaceflight is high delta-v missions, such as missions to the Moon or Mars, can be done more efficiently with more stages. Robert Zubrin has proposed a "mini-Starship" as a 3rd stage for the SuperHeavy/Starship. He trenchantly observed, "The Starship is a reusable Saturn V. It is not a LEM". Addition of a 3-rd stage would allow even higher payload to the Moon. Say, the mass ratio of ~30 to 1 can be retained for the 3rd stage at ca. 300 propellant load and ca. 10 tons dry mass, using a single Raptor vacuum of 380s Isp and 1,900 kN vacuum thrust. This gives an input page result:

 Then the payload to TLI would be ~105 tons:

 This is regarding the SH/SS/mini-SS as a launcher analogous to the 3-stage Saturn V for the Apollo missions. This would require additional in-space stage(s) for doing the landing. However, the payload to TLI is so high this mini-Starship could itself serve as the lander for a round-trip mission to the Moon's surface with a capsule or habitat of ~15 tons mass. This is a mass 50% higher than the mass of the Orion capsule that could be carried not just to a lunar orbit planned for Orion, but all the way to the lunar surface and back to Earth again.

 Flights to Mars.

 Addition of a third stage also allows more efficient flights to Mars in a single launch. This graphic gives the C3 needed for flights to the planets:

 For Mars it is about 14 km2 /s2. The TMI payload to Mars is then:

  A payload of 86 tons sent towards Mars for trans-Mars injection(TMI) as a single flight architecture.

 These are for expendable versions of SuperHeavy/Starship. Elon Musk has estimated the cost of the SH/SS as in the range $100 million to $200 million. Even as expendable this is a greatly reduced cost than the SLS cost of $4 billion per flight.

Estimates of Reusable Payloads.

 The reduction of cost of a flight to the Moon and Mars to only a couple of hundred million dollars rather than multi-billions makes the expendable versions worthwhile. But SpaceX is committed to reusability. 

 Elon has said full reusability would lose 40 to 50% off payload. And even the reusability only for the first stage loses a quite significant 30% from the payload. These are from using powered boost back approach to landing. The propellant that must be kept on reserve, unused during the ascent for the boost back and landing is in the range of 7%. For the 3,400 ton prop load SuperHeavy this ~240 tons. This is doubly disadvantageous for payload in that not only can this prop not be used during flight to orbit but it also acts like additional dead-weight that must be carried in flight.

 Elon Musk has said powered, vertical landing is preferred over horizontal, winged landing because it can be used on airless or low atmospheric worlds such as the Moon and Mars. However, SpaceX needs to do a trade study to see which method results in the least payload lost. Wings typically take up about 5 to 10% of the gross weight of an aircraft. For a vertical launch rocket using non-lifting trajectory on ascent to orbit, this 5 to 10% would only have to be of the dry weight of the craft since aerodynamic lift would be used only during return when the stage is nearly empty. 

 So for the SuperHeavy the added weight of the wings would only have to be 7.5 to 15 tons compared to the 240 tons needed for the vertical landing method. Actually, it likely can be even smaller than the 5%. For one thing rather than using large wings such as used on the Shuttle. It could use short, stubby wings instead. The large wings of the shuttle were due to competing requirements by NASA and the Air Force for where the Shuttle should be able to land. 

Examples of how they could look instead might be those of the X-37 and of the Skylon:

 For the Skylon by using carbon-fiber the wing weight was only 2% of the landed weight. This is 2% of the full gross weight because it used a horizontal liftoff. But since the Starship will be using a vertical liftoff and non-lifting trajectory, the wings only have to support the weight of the vehicle on return, so that 2% only has to be calculated on just the dry weight. The specialty high strength stainless steel used on the Starship might be preferred for the wings as well since it would have about the same strength but have reduced thermal shielding requirements. 

 The landing gear weight can be taken as only 3%:

https://yarchive.net/space/launchers/landing_gear_weight.html

 This likely can be reduced by half to ~1.5% by carbon-fiber or the specialty steel used on the Starship.

 Finally, the thermal protection such as SpaceX’s PICA-X might only add on additional 8% of the dry weight.

  So these extra systems required for reusability will only add a proportionally small amount to the dry mass, and so subtract only a proportionally small amount from the payload.

 

  Bob Clark

Starhopper+Starship as a heavy-lift launcher. Triple-cored Starship for super-heavy lift. 2nd UPDATE, 9/2/2019: Starhopper as a lunar lander.

Copyright 2019 Robert Clark

 Elon has said the current plan is for the BFR first stage, now called the Super Heavy, to have 35 engines, with 6 engines on the Starship for 41 engines total:

e^

@elonmusk

Replying to@fmoflyer

Full stack is 41 rn, but kinda beggin for just one more …

5:17 AM · Jul 21, 2019·Twitter for iPhone

https://twitter.com/elonmusk/status/1152870247612874752?s=20

 It may be possible to accomplish the same payload of a super-heavy lift launcher with a fewer number of engines, and significantly lower cost. But first ...

A Heavy-Lift Launcher.

 A 100 ton launcher is commonly taken as a requirement for manned lunar landing mission. Running the numbers, the BFR’s upper stage in the tanker version, i.e., without the passenger quarters, being used as a first stage booster with an additional StarHopper-sized stage added could form a 100 ton launcher.

 Note there is a tanker version of the BFR upper stage that will not have the passenger quarters and provisions for 100 colonists for a six month flight to Mars, but only an empty fairing. This is the version being discussed here. The Starship version, i.e., the one that does have the passenger quarters, will have significantly greater dry mass than the tanker version.

 The term "Starship" is used in the title only for its current familiarity. It is actually only the tanker version of the upper stage being discussed here.

BFR tanker on left refueling the BFR Starship on the right.

 The tanker version of the upper stage cuts nearly half off the dry mass of the Starship version which has the passenger quarters for 100 colonists on a six month flight to Mars. Then the tanker version would have a dry mass in the range of ca. 45 to 50 tons, at a propellant load of 1,100 tons.

The Starhopper from its size appears to be about in the 400 ton propellant load range. However, the actual Starhopper itself is not weight optimized as it is only intended to make short, low altitude hops. What is needed for the upper stage of this new launcher is a weight optimized stage intended to reach orbit in a TSTO.

Assume we can get this weight optimized Starhopper-sized stage at a ca. 25 to 1 mass ratio, similar to the BFR tanker. Then it would have a dry mass of ca. 16 tons. Then with the 356 s vacuum Isp of the tanker as first stage, and the 382 s vacuum Isp of the Starhopper-sized upper stage, it could get 107 tons to LEO:

356x9.81Ln(1 + 1,100/(50 + 416 + 107)) + 382x9.81Ln(1 + 400/(16 + 107)) = 9,170 m/s, sufficient for orbit.

 SpaceX expects to test launch the BFR’s upper stage, likely in tanker or cargo version, i.e., without the passenger quarters, next year. Since the Starhopper is being built in parallel, SpaceX could probably have the weight-optimized Starhopper-sized additional stage ready in the same time frame. Then you could have a manned lunar mission class launcher by next year, in 2020.

SSTO launcher.

 That the BFR tanker is significantly lighter in dry mass than the Starship version is important. This means the BFR tanker in expendable mode can carry significant payload as an SSTO:

SpaceX BFR tanker as an SSTO.

http://exoscientist.blogspot.com/2017/10/spacex-bfr-tanker-as-ssto.html

A Super-Heavy Launcher.

 The BFR’s 35 engine Super Heavy first stage will likely take longer and be more expensive to develop than the BFR upper stage. I suggest instead that SpaceX develop a triple-cored launcher using the BFR tanker stages as the cores. Judging from the expendable versions of the Falcon Heavy in comparison to the Falcon 9, this could launch about 3 times the payload of the TSTO, so to about 300 tons. This is about what the planned LEO payload of the BFR expendable is expected to be:

e^

@elonmusk

Replying to@joe_mckirdy@13ericralph31 and 2 others

100mT to 125mT for true useful load to useful orbit (eg Starlink mission), including propellant reserves. 150mT for reference payload compared to other rockets. This is in fully reusable config. About double in fully expendable config, which is hopefully never.

2:48 AM · Jul 12, 2019·Twitter for iPhone

https://twitter.com/elonmusk/status/1149571338748616704?s=20

 Elon Musk has given the development cost of the Falcon heavy as $500 million. This is a little more than 50% above the $300 million development cost of the Falcon 9, while being able to launch 3 times as much. 

 Elon on the other hand estimated the development cost of the full BFR as $5 billion. Likely, the triple core version would be significantly cheaper than this. For one thing the triple cores would only take 27 engines plus 3 for the Starhopper upper stage for 30 engines, much fewer than the 41 engines for the planned BFR.

Advantage with Altitude Compensation.
 Another advantage of the triple cores is the increase in payload with altitude compensation. With a typical TSTO, alt.comp. might increase payload 25%. However, with a parallel staged launcher, alt.comp. typically can improve payload 40%. This improves even more so when cross-feed fueling is used in conjunction with alt.comp, typically by 100%. So this triple-cored version with the addition of alt.comp. and cross-feed could have a payload of 600 tons to LEO(!)

Altitude Compensation Improves Payload for All Launchers.
http://exoscientist.blogspot.com/2016/01/altitude-compensation-improves-payload.html


  Bob Clark

UPDATED, 8/13/2019

Another advantage of having a third stage for the BFR at a ca. 250 ton to 300 ton propellant range is that this stage could be launched fully fueled to orbit by the BFR expendable, whether it uses the current plan of a superheavy booster or the triple-cored option. Then a smaller mission size of ca. 25 colonists could be launched to Mars in a single launch, no multiple refueling flights required.

 Judging from the fact the Falcon 9 reusable only reduced a proportionally small amount on the price, this one expendable launch would be cheaper than using 5 to 8 refueling flights. Also based on the long lag time between flights of the Falcon Heavy it could be launched much faster than the refueling, reusable version.

 Note also this small mini-Starship if you will could serve as small SSTO launcher to LEO:

A Small Raptor Spaceship.
https://exoscientist.blogspot.com/2017/10/a-small-raptor-spaceship.html

 I argue this small, reusable SSTO would go a long way toward making spaceflight routine since it would be low cost and could be purchased and operated by independent owners.

 It gets even better. Once orbital propellant depots are in place in LEO, then that one single SSTO once refueled  in orbit can make the full round-trip flight from Earth to the Moon and back again. And if orbital propellant depots are in place at both Earth and Mars then that one single SSTO can make the full round-trip flight from Earth to Mars and back again. 

 Then private, independent owners can make their own manned interplanetary flights.

 See here also for the argument a reusable SSTO can actually be more cost effective than a reusable TSTO because the two-stage loses 50% of its payload on reusability while a SSTO only loses a proportionally small amount:

Case proven: SSTO's are better than two-stage launchers.
https://exoscientist.blogspot.com/2019/08/case-proven-sstos-are-better-than-two.html

UPDATED, 9/2/2019

 The recent successful test hop of the SpaceX Starhopper raises again the possibility of it being used or more accurately a Starhopper-sized stage being used as a lander for the Moon and Mars.

 For brevity, this Starhopper-derived stage weight optimized to have a comparable mass ratio as the Starship cargo/tanker version of ca. 25:1 to 30:1, I'll refer to just as the Starhopper stage. 

For Moon missions, I had originally just wanted the Starship to be used as a first stage and a Starhopper stage to be used as a second stage, i.e., without a Superheavy booster, for a launcher for lunar missions. This would be 100-ton class launcher. It would be cheaper than using the Superheavy booster. In this case, though you would still need a service module propulsive stage and a lander propulsive stage, which would still needed to be designed and constructed.

 NASA has contracted with SpaceX and other space companies for proposals for a lunar lander:

Blue Origin and SpaceX among winners of NASA technology agreements for lunar landers and launch vehicles
by Jeff Foust — July 31, 2019

https://spacenews.com/blue-origin-and-spacex-among-winners-of-nasa-technology-agreements-for-lunar-landers-and-launch-vehicles/

SpaceX also is rapidly progressing with the Raptor engine so that vacuum optimized engines will be available for near term lunar and Mars missions:

SpaceX’s space-optimized Starship engine could be ready sooner than later.
By Eric Ralph Posted on May 23, 2019
SpaceX CEO Elon Musk says that there is now a chance that a vacuum-optimized version of the Raptor engine will be ready for near-term Starship launches, indicating that development has either been re-prioritized or is going more smoothly than expected.

https://www.teslarati.com/spacex-speeds-up-starship-vacuum-engine-development/

 Then if SpaceX were to use a Starhopper stage it would already have the lander for a lunar mission. With a full BFR being ready by 2020, with a Superheavy booster or using triple Starship cores, and the Starhopper only needing to be weight optimized, SpaceX would have a full lunar landing capable rocket already in 2020.

 Moreover, with the Starhopper stage being able to be delivered to LEO fully-fueled by the BFR in expendable mode, the Starhopper would then have sufficient capability to carry the Dragon 2 capsule from LEO to the Moon's surface and back to Earth in a single stage. No refueling flights required.

 In fact, it would have the capability to carry the Orion capsule and its service module at ca. 16 tons dry mass to the Moon and back, as long as the service module was unfueled, with all propulsion at the Moon being done by the Starhopper stage.

 Both the Dragon and Orion capsules are rather small in diameter compared to the 9 meter diameter of the BFR though. You would need a rather large adapter for either of them. Then instead we could use a large hab for the purpose. Being able to deliver and return ca. 16 tons to the lunar surface, we could use a Transhab-sized crew/passeger module.

 The Transhab was designed to be a habitat for a several months long Mars mission. It was to carry a crew of 6 at a mass of 13 tons and a 340 cubic meter volume. The Transhab was inflatable but at an inflated diameter of 8.2 meters it could be launched fully inflated on the BFR.

https://en.wikipedia.org/wiki/TransHab 

 The Transhab could be transported to the lunar surface for longer stays a la the 6 month crew rotations on the ISS. The Starhopper would have enough capability so that the same stage could also return it to Earth without needing refueling.

 Additionally, the Starhopper could deliver 35 tons of cargo to the Moon in a reusable mode where it returned to Earth after dropping off the cargo, or 55 tons to the Moon in a one-way, expendable mode. This would go a long way towards constructing the Moon base.


 Being able to deliver 55 tons to the lunar surface becomes quite important when you consider the size of the Starship's passenger quarters. Since the Starship with the passenger quarters masses 85 tons, but the cargo/tanker version without the passenger quarters masses at 45 to 50 tons, we can estimate the Starship passenger quarters for 100 colonists on a months long flight to Mars as 35 to 40 tons. Then this can be transported to the Moon with 100 passengers for long stays on the Moon by the Starhopper lander. In fact, it could form the basis for a lunar base, or colony.

 Image of the Starship with passenger quarters. For the Starhopper version the tankage section would be 1/4th the size.