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

Possibilities for a single launch architecture of the Artemis missions, Page 3: Saving the lander mission for Artemis III.

 Copyright 2023 Robert Clark

  I discussed a possible single-launch lunar lander architecture here:

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

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

 The SpaceX delay in the Starship HLS development has led to NASA considering that Artemis might not even be a lander mission. This leaves open the possibility to save the lander mission for the Artemis III mission alternative methods for landers should be considered.

 Plus, there is the fact many knowledgeable space aficionados from the old days really do not like the SpaceX plan of using 8 to 16 refueling flights just for one lunar lander mission.

 This plan for a replacement lander could be done rather quickly and at low cost because it would use already existing space assets. Also, it would be done by our European partners so would not require NASA expenditures using all European space components. That would save $3 billion that NASA would have had to pay to SpaceX for their lander.

 This would involve even greater European involvement in successfully accomplishing the Artemis missions than just Orion’s service module so would undoubtedly get enthusiastic support from the ESA.

 It is notable that the ESA has expressed even greater support for lunar colonization plans than even NASA. This ESA produced lunar lander would allow them to further their own plans for a sustained human presence on the Moon.

 About ESA’s ATV-derived service module for Orion, that again would require low cost modifications in this plan. It would need just an addition 10 tons of propellant, which would fit easily within a service module diameter expanded to match the Orion’s diameter. Again this cost would be covered by our European partners, with no expenditure by NASA.

 Those two factors would be the easiest aspects of the plan. It might be difficult to believe a lunar lander would be among the “easiest” parts of the plan. But keep it mind it would be derived from already existing space assets.

 The trickiest aspects of the plan would be the fact the SLS would require higher payload capability to allow for the higher propellant load of the service module of 10 tons and a ca. 15 ton mass lunar lander.

 One possibility, keeping the Boeing EUS, is to put atop it a third stage consisting of the 50-ton Centaur V, as discussed in, "Possibilities for a single launch architecture of the Artemis missions, Page 2: using the Boeing Exploration Upper Stage."

 However,  I’m still not convinced the Boeing EUS is the best way to go because of its expense and its small size. If a cryogenic upper stage at a propellant load of ca. 200 ton size instead of the Boeing EUS ca. 125 tons were used, then this larger upper stage itself could do TLI burn carrying the Orion, larger SM, and ca. 15 tons lunar lander.

 I discuss here how such a larger cryogenic stage could be done in a much cheaper fashion than the Boeing EUS approach:  

Why does the Boeing Exploration Upper Stage(EUS) cost so much?  

http://exoscientist.blogspot.com/2022/11/why-does-boeing-exploration-upper.html

 I call this the trickiest aspect of the plan because Boeing has shown repeated delays in getting the core stage ready so the same might happen with their EUS stage, especially when it would have to be moved up to be ready by the 2025 Artemis III launch date, instead of on Artemis IV in 2028. As for the extra Centaur V third stage, since it is expected to first launch this year, likely it will have several launches under its belt by a 2025 Artemis III launch date.

 Note though a MAJOR reason why the development of the different versions of the SLS was arranged as it was was because of cost reasons. The development of the Boeing EUS was pushed back to delay paying for its wildly overpriced development costs. Note too not having to pay for the SpaceX Starship lander would save NASA $3 billion.

 Boeing’s charge to NASA for the EUS is a key reason why I prefer the simpler approach for an upper stage of just basing it on the core with fewer barrel rings. 

This would also give us the stage more cheaply and more quickly since it involves just using fewer rings on the tanks. If you have ever watched the video of construction taking place at the SpaceX development site, barrel rings of the tanks on the Starship and SuperHeavy are swapped out, replaced, taken-off and put back routinely.

 You’ve heard the mantra of former NASA administrator Dan Goldin, “faster, better, cheaper”? This would be "faster, better, cheaper, and simpler".

 Since this is of different design though that also brings into question its availability by a 2025 launch date. Note that this plan involves several new components. For that reason, we might want to use Artemis III as an unmanned test lander mission. We might even have it be “manned” by human-like robots, with their operation controlled from the ground on Earth.

 In this regard it is notable that several lines of evidence suggest that there might be valuable metals at the lunar South Pole, the planned location for the Artemis III landing. Indeed, the untold trillions of dollars of valuable metals speculated to exist in the main-asteroid belt on 16 Psyche might already exist just next door at the Moon's south pole! I discuss this here:

U.S. will lag behind in utilization of resources on the Moon.

http://exoscientist.blogspot.com/2023/08/us-will-lag-behind-in-utilization-of.html

 By the way, the title there stems from my dismay that the U.S. rovers to the South Polar location won’t have instruments for detecting heavy metals but the rovers from other countries will. This is such an obvious thing to include, especially when other countries will include them, that it’s mystifying why the U.S. chose not to include them.

 In any case, it would be pretty cool seeing human-like robotic astronauts prospecting for valuable metals at their landing site.

The ESA produced lunar lander I suggest using is of Apollo-like size at ca. 15 tons. But its crew module volume would be much larger as it is based on the Cygnus capsule, given life support. The Apollo lunar lander had a 6.7 cubic meter internal volume. But the Cygnus has an 18.9 internal volume, nearly 3 times that of the Apollo LEM and the expanded version of the Cygnus has a 27 cubic meter internal volume.

 But as for sending cargo or habitats to the Moon, the SLS is far too expensive, $2 billion+ per launch, and at too low flight cadence, at best 1 once per year, for that purpose. 

 Better to use lower cost launchers such as the Falcon Heavy for the purpose. I estimate using all hydrolox in-space stages, given low-boiloff tech, the FH could get 15 tons one-way to the lunar surface. 

 Using hydrolox only for the TLI burn but a storable propellant lander stage (so no low-boiloff tech needed), the FH could get 10 tons to the lunar surface.

Robert Clark

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

 Copyright 2023 Robert Clark

  A comparison between the Apollo and Orion capsules:

 Rarely has a design mistake been so clearly illuminated by a single picture. Note the Orion capsule is nearly double the size of the Apollo capsule in mass. But rather than making Orion’s Service Module twice as big as the Apollo Service Module, as it should be to get similar performance, instead it is smaller by 1/3rd.

 Orion’s service module is based on ESA’s ATV cargo tug to the ISS, which had a 4.5 meter diameter and a 10 ton propellant load.

BUT THERE WAS NO REASON TO KEEP IT AT THAT SAME DIAMETER FOR THE ORION USE, NOR TO KEEP THE SAME SIZE PROPELLANT LOAD.

 If instead the diameter was made to match the capsule’s diameter, as was the case with Apollo, there would be an additional 20 cubic meters of volume inside the Service Module, well more than enough to hold an additional 10 tons of the storable propellant used.

And that is all that is needed to solve THE major problem of the SLS/Orion approach: the fact it can’t send the Orion and a lunar lander to low lunar orbit, and bring the Orion back to Earth again.

 It is because of that the idea of the lunar Gateway was proposed, where the SLS would only have to take the Orion to a further out orbit.

 But if instead the Service Module was given that additional 10 tons of propellant then it could send both the Orion and a ca. 15 ton lunar lander to low lunar orbit, and have enough propellant left over to bring the Orion back to Earth, a la the Apollo architecture.

 Rarely, has a mistake been so clearly exposed, especially when its solution is so clearly made apparent as well.

 In the blog posts, "ESA Needs to Save NASA's Moon Plans", and "Possibilities for a single launch architecture of the Artemis missions", I wrote about getting a single launch format for the Artemis lunar lander missions by using the Ariane 5 as an upper stage or by using two Centaur V stages as the upper stage for the SLS, respectively.

 This stemmed from dislike of the plan NASA was endorsing of using multiple flights and refuelings of the SpaceX Starship as the lander. I also objected to the high cost projected for the planned Boeing Exploration Upper Stage(EUS), being nearly half the cost of the entire SLS per flight, nearly $1 billion.

 However, NASA has negotiated a better price structure for the EUS. And it appears NASA is wedded to the Boeing EUS. Then I'll discuss a single launch architecture using the Boeing EUS upper stage.

 The payload to LEO of this version of the SLS with the Boeing EUS, which is version Block 1B, will be 105 tons to LEO. The current fueled mass of the Orion+Service Module is 26.5 tons. An additional 10 tons of propellant will bring it to 36.5 tons. 

 In the blog post, "A low cost, lightweight lunar lander", I discussed a lunar lander at a 13-ton total fueled mass based on the Cygnus capsule given life support as the crew module, and the Ariane 5 EPS storable propellant stage as the propulsive stage for the lander.

Calculations for the delta-v to the Moon and back.

 The Orion with its fully fueled service module has a mass of 26.5 tons. The propellant load of the service module is ~10 tons, with 16.5 total tons dry mass of the Orion and service module. We'll add an additional 10 tons propellant to the service module to bring the total mass to 36.5 tons, including 20 tons of propellant.

 The AJ-10 engine used has a vacuum ISP of 319s. We'll assume a fueled lunar lander of size ~13 tons, as described in the blog post, "A low cost, lightweight lunar lander", comparable in size to the Apollo missions lunar lander. So, a 16.5 + 13 = 29.5 ton mass for the vehicles that need to be put in low lunar orbit. But remember also we need to have some propellant left over in the service module to bring the Orion back home to Earth.

 For the delta-v calculation, after the SLS places the Orion/Service Module/lunar lander stack in trans-lunar injection(TLI) towards the Moon, we need 0.9 km/s to put the stack into low lunar orbit. This requires 13 tons of propellant, leaving 7 tons remaining:

319*9.81Ln(1 + 13/(29.5 +7)) = 0.950 km/s. The lunar lander will then be launched to land on the Moon while the Orion and service module remain in lunar orbit.

 After the lander mission is completed, the lander returns the astronauts to the Orion in lunar orbit, and the lander is then jettisoned. The Orion's service module is then fired to bring the Orion back to Earth. After lander jettison, the dry mass of the Orion and service module will be 16.5 tons. Then the 7 tons of remaining propellant is sufficient to perform the trans-Earth injection(TEI) burn of 900 m/s to escape lunar orbit and place the spacecraft back onto the free return trajectory back to Earth:

319*9.81Ln(1 + 7/16.5) = 1,100 m/s. 

 So the total mass that needs to be sent to trans lunar injection(TLI) on a path to encounter the Moon is 36.5 + 13 = 49.5 tons. Now use the rule-of-thumb that a Centaur-like hydrolox stage can send to TLI at a 3,000 m/s required delta-v a payload mass equal to its propellant load.

 So use for a third stage atop the Boeing EUS the Centaur V at an 50 ton propellant load and 5 ton dry mass. This then results in a total mass to LEO of 104.5 tons consisting of the 55 tons of the Centaur V plus the 49.5 tons of the Orion capsule/Service Module/lunar lander, within the lift capacity of the SLS Block 1B to LEO.

  Robert Clark

Possibilities for a single launch architecture of the Artemis missions.

 Copyright 2022 Robert Clark

 In the blog post ESA Needs to Save NASA's Moon Plans I noted that the original plan SpaceX submitted to NASA for a lunar lander required 16 launches due to multiple refueling flights, with the refueling flights to orbit requiring a time of 6 months to accomplish. I argued in the blog that if instead NASA used an Ariane 5/6 as the upper stage of the SLS rocket replacing the current Interim Cryogenic Propulsion Stage(ICPS) then it could be done in just a single launch of the SLS, with no launches of the Starship required at all.

 After their proposal was submitted by SpaceX and accepted by NASA, Elon Musk, stung by the criticism it would take so many launches, suggested it probably could be done in only 4 refuelings since a stripped down Starship for a lunar lander mission would weigh much less.

 SpaceX needs to be open about what the mass would be for such a stripped down Starship since that would directly affect how much NASA, and the U.S. taxpayers, would have to pay to SpaceX for refueling launches. See discussion here, 

The nature of the true dry mass of the Starship. 

 My suggestion to use the Ariane 5/6 as an SLS upper stage was critiqued on political acceptability grounds for a such a large contract to be taken from a U.S. company and given to a European company. 

 Here I'll propose a solution using existing, pretty much, American upper stages for the SLS. It's the ULA Centaur V upper stage coming into service next year. I considered using the Delta IV common core stage but at a 40 meter height it might be too tall for this use.

 

Architecture.

 The Centaur V has a 54 ton propellant load. Following the approx. 10 to 1 gross mass to dry mass ratio of the original Centaur, I'll take the dry mass to be ~5 tons. Then I'll examine  two options: 1.)2 Centaur V's combined into a single stage, and 2.)2 separate Centaur V's.

 The current Block 1 version of the SLS gets about 27 tons to trans-lunar injection(TLI). This is the speed needed to get a spacecraft once in orbit to reach the Moon. The 27 tons is just enough to get the Orion capsule and its service module to TLI. 

 However, the current approach is not to put the Orion in low lunar orbit around the Moon. Instead, it will be placed in a higher altitude orbit of Earth-lunar space called a near-rectilinear halo orbit(NRHO). The reason is the current version of the SLS did not have enough power to put the Orion in low lunar orbit and for it to be able to escape again.

 Our plan then is to first increase the payload capacity of the SLS so that enough additional propellant can be given the Orion service module so the Orion can actually reach and leave low lunar orbit. 

 The Orion with its fully fueled service module has a mass of 26.5 tons. The propellant load of the service module is ~10 tons, with 16.5 total tons dry mass of the Orion and service module. We'll add an additional 10 tons propellant to the service module to bring the total mass to 36.5 tons, including 20 tons of propellant.

 The AJ-10 engine used has a vacuum ISP of 319s. We'll assume a lunar lander of size ~15 tons, comparable in size to the Apollo missions lunar lander. In a following blog post we'll describe it in more detail. So, 16.5 + 15 = 31.5 tons dry mass needs to be put in low lunar orbit.

 For the delta-v calculation, after the SLS places the Orion/Service Module/lunar lander stack in trans-lunar injection(TLI) towards the Moon, we need .9 km/s to put the stack into low lunar orbit. This requires 13 tons of propellant, leaving 7 tons remaining:

319*9.81Ln(1 + 13/(31.5 +7)) = .910 km/s. The lunar lander will then be launched to land on the Moon while the Orion and service module remain in lunar orbit.

 After the lander mission is completed, the lander returns the astronauts to the Orion in lunar orbit, and the lander is then jettisoned. The Orion's service module is then fired to bring the Orion back to Earth. After lander jettison, the dry mass of the Orion and service module will be 16.5 tons. Then the 7 tons of remaining propellant is sufficient to perform the trans-Earth injection(TEI) burn of 900 m/s to escape lunar orbit and place the spacecraft back onto the free return trajectory back to Earth:

319*9.81Ln(1 + 7/16.5) = 1,100 m/s.

Calculations for Earth escape stage to TLI.

 That's the plan if we can upgrade the SLS to carry sufficient payload to give the Orion service module that extra 10 tons of propellant. The total mass that needs to be put into TLI is 36.5 + 15 = 51.5 tons. Here's a calculation for the first approach of two Centaur V's combined into a single stage. I'll use the payload performance calculator of Dr. John Schilling, on Silverbirdastronautics.com. The specifications for the 5-segment SRB's are taken by scaling up the numbers from the 4-segment SRB's used on the Space Shuttle system.

 I'll give this stage 4 RL10 engines instead of the Centaur V's 2 because of the larger size, in effect just transferring two of the RL10's from the second Centaur's to the first. The input page looks like this:

                                                               

 The payload estimator then gives the payload to LEO of ~127 tons:

 

  And the for the payload to TLI we'll use a C3 of -1.00km2/s2.

 This gives a payload to TLI of about ~52 tons:

  It is notable though the Schilling payload estimator has rather large error bars. These numbers need to be confirmed by more accurate payload estimators.

 The payload can be increased by using instead of the RL10's, a single Blue Origin BE-3U, the vacuum optimized version of the BE-3 engine used on the New Shepard. This engine has a vacuum optimized thrust of 710 kilonewtons. Placing this in for the upper stage thrust gives a payload to LEO of 136 tons, and to TLI of 54.7 tons. Again this needs to be confirmed by more accurate payload calculators.

 The intent here is to find a low cost approach to an upper stage that would allow a single launch architecture for the Artemis lunar lander missions. A combination of adding additional engines and also combining two tanks would ratchet up the costs.

  The second approach would use two separate Centaur V's. However, because of the large mass that needs to be carried by the either Centaur as payload we'll give both Centaurs 4 RL10's. The input screen looks like this on the Schilling calculator:

  And the LEO payload is ~129 tons:

 And the TLI payload is ~54.7 tons:

  Again, these payload estimates would have to be confirmed by more accurate payload estimators.

 This second approach would not incur the extra costs of combining two Centaur V's into a single stage, but it would require 4 additional RL10's. As before though we could get increased payload by replacing the RL10's by the BE-3U, and likely lower cost.

 We still need to come up with that lunar lander of comparable gross mass as the Apollo lander, ~15 tons. In a following blog post I'll show our European partners can come up with such a lander at low cost and at a relatively short time frame.

  Robert 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.