SpaceX routine orbital passenger flights imminent.

 Copyright 2024 Robert Clark

 An approximate $100 per kilo cost has been taken as a cost of space access that will open up the space frontier. For then instead of a price for a private citizen to go to space instead of millions of dollars, it could be priced at a few tens of thousands of dollars. This is in the range of the price of a first-class roundtrip ticket from Los Angeles to Australia. SpaceX now has the capability to offer such a launcher at such a low per kilo rate.

Robert Zubrin has said in an interview that Elon Musk informed him he believes he can build the Starship, i.e., the upper stage of the Superheavy/Starship launcher, for $10 million:

SpaceWatch.Global on X: "“We don’t go to Mars to desert the Earth. We go to Mars to expand the capacity of the human race, to create new branches of human civilization.” - Dr. Robert Zubrin In this insightful Space Café Podcast episode, Dr. @robert_zubrin dives into the real challenges of building https://t.co/A41iFzClpY" / X

 But Zubrin notes what SpaceX is aiming for is reusability. Say a new Starship might cost $10 million to build with a purchase price of $20 million to the customer. Then allowing conservatively 10 reuses SpaceX might only charge $2 million per use. However, he does not mention it here but he is implying the use of this as a launcher independent of the first stage Superheavy. Then you would need a smaller upper stage, a mini-Starship. 

 Zubrin has discussed use of a mini-Starship, but in the context of a 3rd stage for the Superheavy/Starship. Presumably here though, he is suggesting a smaller launch system consisting of the Starship now as a first stage and a mini-Starship as an upper stage.

 An upper stage is commonly 1/4th to 1/3rd the size of the previous stage. So call the cost of the mini-Starship new of, say, $3 million. However, in regards to reusability, SpaceX has found that difficult to implement for an upper stage, particularly in regards to the thermal protection system. Afterthe last test flight IFT-6 for example, Elon Musk has suggested they might have to change to a completely different kind of TPS than the ceramic tiles now used. In contrast though, SpaceX has ampy demonstated with the Falcon 9 booster the first stage is much easier to reuse.  So I'll estimate a cost here of a partially reusable Starship/mini-Starship as $5 million.

 We'll calculate here that this smaller Starship/mini-Starship launcher will still be a a Saturn V-class expendable launcher at 100+ ton payload capacity to LEO. It comes from this Elon Musk estimate of the dry mass of the Starship as an expendable:

Elon Musk

@elonmusk

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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.

9:14 PM · Mar 29, 2019

 But that is for an upper stage use where it did not have enough engines for liftoff from ground. Assume for 1st stage use it needs 9 engines. Increase the dry mass now to 50 tons for the greater engine mass. 

 For the mini-Starship, an upper stage commonly is 1/3rd to 1/4th the size of the lower stage, so call it 420 tons propellant mass. As an upper stage it doesn’t need high engine thrust so assume same mass ratio of ~30 to 1 as for Elon’s expendable Starship estimate, giving it a dry mass of 14 tons.

Take Starship exhaust velocity as ground launched as comparable to that of the Superheavy, 3,500 m/s. And take the upper stage’s vacuum exhaust velocity as 3,800 m/s. Then we could get ~120 tons to LEO: 

3,500Ln(1 + 1,200/(50 + 434 + 120)) + 3,800Ln(1 + 420/(14 + 120)) = 9,200 m/s, sufficient for launch to LEO.

 This a price of only $5 million launch cost for a Saturn V-class launcher as partially reusable. If we make the landing downrange we lose only ~20% off the expendable payload judging by the Falcon 9 example, so still ~100 tons to LEO as partially reusable. That amounts to radical reduction in launch cost down to only ~$50 per kilo. This means that rather than a price to orbit for a passenger being millions of dollars as it is now, it could be in the tens of thousands of dollars range. As noted by Zubrin in that SpaceWatch.Global interview this is in the range of a first-class round trip tocket from Los Angeles to Australia.

 Those price estimates though are based on a $10 million cost of the Starship. But that undoubtedly is for high production rates. 

 So we'll use an estimate based on the current production cost for Superheavy/Starship at about $90 million, with about 30%, $27 million, for Starship:

STARSHIP COST ANALYSIS
OVERVIEW

Note: This is Payload's current estimate and not based on access to any internal Space data or proprietary information.

Current Estimated Starship & Booster Full Stack Cost (S in thousands)

39 Raptor Engines                                   39,000

Labor.                                                       35,000

Structure, plumbing, tiles, parts               13,000

Avionics                                                     3,000

Total                                                         90,000

*Payload costs estimates are based on a post-R&D 1-2 year forward-looking model. This is an educated best estimate and not based on Space internal data. Further cost reductions are expected in the long-run. $90M cost: Payload estimates it costs $90M to manufacture a fully integrated Starship based on a post-R&D/test production phase near-term model. The go-forward cost does not factor in the near $5B SpaceX has spent on R&D to date.

~70% of costs accrue to Super Heavy and ~30% to Starship upper stage.

Future Starship (upper stage) cost reductions: As Starfactory comes online and Raptor production is refined, Space aims to reduce costs even further. A focus on Starship's upper stage: When SpaceX achieves full reusability, production of Starship second stage vehicles will be an order of magnitude higher than booster production.

• The company plans to eventually build multiple second stage Starships per week and reduce

Raptor engine's production cost to $250K a pop. If successful, the long-term cost to mass produce second-stage Starships could drop to $10M to $15M a vehicle. However, for purposes of this report, we will analyze costs as they are today.

Raptor 2 engines ($39M) Payload estimates each Raptor 2 engine costs ~$1M to build. The 39 engines-which include three additional upper-stage engines that will be added in the future-are by far the biggest Starship cost, adding $39M to total cost. SIM per Raptor 2 engine is half as expensive as its $2M+ Raptor 1 predecessor. 20 SpaceX hopes to eventually bring the cost per engine down to ~$250K.

Payload Research
18. Elon Musk on X 19. Space 20.Elon Musk on X 21. Elon Musk on X

https://docsend.com/view/fi9wuazzeex57iig

 Say, for a mini-Starship upper stage its cost would be a 3rd of the $27 million current production cost of the Starship as new, so $9 million. So the full vehicle production cost at $36 million as new. So a Saturn V-class launcher capable of 100+ tons to LEO at ca. $36 million price new. Note this is about half that of the price of the Falcon 9 but at 5 times the payload capability. This is a cut in price per kilo by a factor of 10 down to $300 per kilo from $3,000 per kilo.

 Note again though with SpaceX amply demonstrating practicality of first stage reusability we can do better than this still. Say the Starship now as first stage could be reused 10 times cutting its cost to, say, $2.7 million per launch, for the total partial reuse cost of $11.7 million per launch. But this is the price to SpaceX. Double this for a price to the customer of ~$23 million as partially reusable. As before landing downrange for the booster would still allow ~100 ton payload to LEO, for a price per kilo of $230 per kilo.

 This may still allow passenger tickets to orbit at say the hundreds of thouasands range depending on how many passengers could be carried in a passenger cabin. Note this is price range charged by Virgin Galactic and Blue Origin on New Shepard just going to suborbital space. Most importantly, large numbers of launches carrying private passengers to orbit of wealthier customers at the price point doable now will increase the production numbers for the Starship thus enabling the lower price point to be reached. 

Launch costs for manned Moon or Mars flights at only ~$20 million per launch.
 This is a Saturn V-class vehicle capable of single launch Mars or Moon missions we could launch now. No thermal tile problems, or needing to master orbital refueling, or stretching tanks, or increasing Raptor thrust. It literally could have been launched on the last few test launches and can literally be launched on the next test launch, providing a proof-of-principle for manned flights to the Moon or Mars. Note this is less than the cost now for sending astronauts to the ISS.

 I argue that this is better than the currently planned SpaceX/NASA approach. For instance the fully reusable Superheavy/Starship V2 will have payload at 100+ tons, and still need all of orbit capable TPS, orbital refueling, tank stretch and upgraded raptors. And it will need all of these to get a ca. $10 million reusable launch cost. But the Starship/mini-Starship will reach the same payload capability, at only be 1/3rd the size of the Superheavy/Starship, without the difficult technical advances.

The current approach to the Moon is not just worse, it's multiply times worse. SpaceX wants a multiple refueling, multiple launch approach of the full Superheavy/Starship for lunar missions. This will be in the range of 18 total flights with refuelings, the orbital depot, and the Starship HLS itself.

 In contrast the Starship/mini-Starship can do it in a single launch. And the current plan needing ~18 launches will actually be 18*4 = 72 times bigger than using a single Starship. Taking into account the size also of the mini-Starship, the current multiple Superheavy/Starship lunar plan will be about 50 times the size of just the single Starship/mini-Starship.

 Using an existing Falcon 9 upper stage or Centaur V as the 3rd stage/lander this is a capability we have now to do single launch Moon or Mars missions. 

 Keep in mind the ~$20 million price might be the customer price for the launch-to-LEO vehicle of the Moon or Mars flight. But SpaceX's own cost would be ~$10 million per launch. This is a capability SpaceX has now. 

 A comparison of the relatively sizes of the two approaches for getting to the Moon:

 Compared to:

  

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