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Everyday Astronaut suggests a single launch architecture for a Moon mission using an expendable Starship as the launcher, so eliminating the SLS, and using Blue Moon as the lander, and the Orion as the capsule: 

Everyday Astronaut @Erdayastronaut
Call me pessimistic, I honestly think the only way to get humans back on the moon in this decade is likely a single launch approach, but not with SLS. SLS's TLI capacity is insufficient. Honestly, make an expendable Starship upper-stage with a Saturn V style clamshell stage adapter to house a fully fueled Blue Moon Lunar Lander inside and an Orion Capsule on top.

Booster might even be able to be caught and still get all that mass on a TLI with a dedicated expendable upper stage. If not, just ditch the booster for maximum performance, who cares, that's a drop in the bucket compared to even a single RS-25. It could literally be demonstrated in less than one year from today on the Starship side.

Without needing a refueling architecture from Blue Moon, it also simplifies that development process. A stage adapter for Orion is very simple.

Make no mistake, orbital refueling is the future and will be what makes these things more sustainable, but if we're in a race to beat China to the moon, the first return to the moon could be done on a very accelerated timeline without relying on such lofty milestones and infrastructure. 

Maybe I'll do a video breaking down this and other options when I finish my other video.
https://x.com/Erdayastronaut/status/1965141937931120910

  Bob Clark

It would be like spending all the money they did on the Falcon 9 and winding up with a rocket of less payload than the Falcon 1. We would say they took a bad engineering approach to the development of the larger rocket.

  Bob Clark

I'd like to see SpaceX do experiments using actual wings for the return, not just the "flaps" now used. If the wings are sufficiently lightweight Starship might need no thermal protection at all:

Wings in space.
by James C. McLane III
Monday, July 11, 2011
Wing loading (the vehicle’s weight divided by its wing surface area) is a prime parameter affecting flight. The antique aluminum Douglas DC-3 airliner had a big wing with a low loading of about 25 psf (pounds per square foot of wing surface). At the other end of the spectrum, the Space Shuttle orbiter has a high wing loading of about 120 psf. This loading, combined with an inefficient delta-shaped wing, makes the orbiter glide like a brick. A little Cessna 152 private plane features a wing loading of about 11 psf and modern gliders operate down around 7 psf. A space plane with huge lifting surfaces and a very low wing loading might not require any external thermal insulation at all. Building a space plane with a wing loading of, say, 10 psf should not be an impossible proposition. Perhaps some day it will be done.
http://www.thespacereview.com/article/1880/1

  Bob Clark

Time for the polls again about the success of SuperHeavy and Starship for the upcoming launch:

About SuperHeavy:

1.) SuperHeavy will ascend with no engines out and do boostback and landing burns with no engines out for a soft ocean landing.

2.) SH will have 1 or more engines out during ascent and/or the return burns but still complete the ocean touchdown.

3.) SH will explode either during ascent or during return burns.

About Starship:

1.)Starship will complete its burns with no engine outs.

2.)ST will have 1 or more engine outs either during ascent or return burns but still complete ocean touchdown.

3.)ST will explode either during ascent or return burns.

Bob Clark

It had been thought that hydrolox was the best propellant to use for a rocket SSTO because of its high ISP. But closer analysis revealed that dense propellants are better for the purpose because their higher density results in lower tank mass, a significant component of the vehicle dry mass.

See discussion here:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
https://forum.nasaspaceflight.com/index … ach=587468

The tankage for hydrolox is about 3 times heavier, based on density, than kerolox for a rocket SSTO. But for an airbreather it’s even worse: about 10 times heavier!

  Bob Clark

Skylon Triumphant! Towards the SSTO.

The new program just announced called Invictus to continue the Skylon project will use a standard jet-fuel engine instead of hydrogen-fueled though it will still be precooled using hydrogen.

This may be an unrecognized blessing. My opinion is the major stumbling block from Skylon progressing is it assumed a hydrogen-fueled jet engine and such did not and still does not exist.

Nothing beats having that engine and precooler right in front of you where you can vary engine and precooler parameters to optimize performance. Skylon instead had to simulate this since they did not have a hydrogen engine at hand.

By taking the approach of using an existing jet-fuel engine, the American hypersonic concern Hermeus is rapidly proceeding to actual test flights. I predict the same will be possible with Invictus.

However, a question I had involved an aspect of the graphic attached below showing the T/W ratio and specific impulse of the Skylon compared to other propulsion methods.

The notable aspect of the Skylon Sabre engines is they used a precooler ahead of the turbojet engines. But what’s so remarkable is it increases the thrust by a factor of 3. The specific impulse is actually reduced(though still high as an airbreather) but that tripling of the thrust for an airbreathing engine is surprising.

The importance of this is that it has been frequently argued about producing an SSTO of why not just put some jet engines on it for the first part of the flight? The problem is jet engines are so heavy. Commonly their T/W ratio is only ca. 5 to 7.

The key question arises from the fact that graphic shows with a precooler the T/W ratio of a hydrogen airbreathing engine can be increased by a factor of 3. Will that increase also apply for a jet-fueled turbojet for the thrust to be increased 3 times?

The highest T/W ratio jet engine we have is at ~11.5 in the F135 engine. Then with precooling could the T/W ratio be increased to the range of 35 to 1??? If so, then that would be a radical advancement in the feasibility of an SSTO, since the weight of the jet-engines will be approaching that of rockets while still having the high Isp of airbreathers.

  Bob Clark

I do think a combined-cycle airbreathing+rocket vehicle can work as an SSTO, however there is a key gap in your argument. Looking at the article you cite:

Binbin Lin, Hongliang Pan*, Lei Shi and Jinying Ye
Effect of Primary Rocket Jet on Thermodynamic
Cycle of RBCC in Ejector Mode
Int J Turbo Jet Eng 2017; aop
https://drive.google.com/file/d/16jeWe2 … hP__9/view

it requires scramjet propulsion. But no one has gotten a scramjet to work for more than a few seconds of positive net thrust.

Perhaps you should aim first for ramjet only needed for the upper limits of the airbreathing part of the flight to ca. Mach 5 to 6.

This is the approach aimed for by Skylon. Skylon has been reborn by the way:

INVICTUS: Europe’s Advanced Hypersonic Test Platform by ESA/Frazer-Nash.
https://www.youtube.com/watch?v=zwMHvsNCmQE

I am inviting contributions to my LinkedIn group on SSTO here:

SSTO - Single Stage to Orbit.
https://www.linkedin.com/groups/13205030/

  Bob Clark

About injection fuel or oxidizer or even inert gas like nitrogen horizontally from around the sides of the throat, I’m not intending this to burn. It is simply to constrict the size of the area the already combusted gases can pass though to the nozzle. It would be at comparable pressure or higher to how the propellant is injected into the combustion chamber from the feed lines.

  Bob Clark

Thanks, GW for the discussion of getting the Isp vs. altitude with adaptive nozzles. I’ll give it a try to see if I can get it to work. I suppose I can test my calculations by first doing it for the Vulcain engine and seeing if I get the same results as that graphic.

In regards, to the extended nozzle being too large for the rocket diameter, there are other ways of doing the altitude compensation. For instance what’s key is the expansion ratio has to vary. This can be done by keeping the nozzle fixed but varying the nozzle throat. This was done for the ASALM ramjet missile for example.

  Bob Clark

Yes. We know ramjet propulsion is doable as operational ramjets have been fielded since decades ago. What we haven’t been able to do is scramjets that can work for more than just a few seconds of positive net thrust. But ramjets can operate up to Mach 5+, over 1.600+ m/s. This can subtract off a significant amount of the delta-v needed for orbit. This can certainly work for TSTO where the first stage is airbreathing ramjet, and the upper stage being a rocket. This can cut the cost to space since the airbreathing first stage might be reusable for thousands or reuses.

It might also work for a airbreathing/rocket combined-cycle SSTO. I am investigating this possibility.

   Bob Clark

Repeated engineering failures stem from the top. An analogy, suppose a wealthy businessman started his own civil engineering company and named himself Chief Engineer, despite his background not being engineering.

OK, it’s his company he can name himself anything he wants. But suppose as Chief Engineer he then proceeds to ignore basic principles of civil engineering. Would anyone be surprised if his buildings and bridges fell down?

Why SpaceX needs a True Chief Engineer.
http://exoscientist.blogspot.com/2025/0 … ineer.html

Bob Clark

I did a search of “RGClark” on the forum, to find posts I hadn’t seen and saw this. I don’t remember it. Perhaps you mean Sidereus that want’s to make a small SSTO:

Sidereus Space Dynamics Complete Integrated Static Fire Test.
https://europeanspaceflight.com/sidereu … fire-test/

  Bob Clark