Alternative Shuttle Development Universe

[Scott Kenny]

Splitting hairs there - transferring propellant from one system element to another *IS* crossfeeding.

And I'd argue that feeding LOx and LH2 from the external tank to the engines on the orbiter is definitely crossfeeding.


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If we're positing a system that isn't using SRBs, I can make a good argument for tripropellant. KeroLOX at initial launch, feeding LH2 in as you climb up. You need the sheer mass flow of kerolox if you aren't using solids at liftoff.

 

[martinbayer]

And I'd argue that feeding LOx and LH2 from the external tank to the engines on the orbiter is definitely crossfeeding.

That's exactly where I'm coming from - the ET and Orbiter simply were two separate and different system elements!!!

 

[publiusr]

Now if an upper stage is methalox but small, and a winged booster is tri-propellant---I guess what I am trying to do is gin up a system where an expendable upper stage really is super expendable (besides the payload). Folks in the 70's may have felt similarly.

The fly-back booster nursemaids the upper stage, so what de-orbits is little more than pot-metal...no SSME mass power heads coming in at Mach 2 like Columbia's break up.

Don't know about Starship engines.

Seeing the Merlin upper stage nozzles expended irks me.

 

[Scott Kenny]

Now if an upper stage is methalox but small, and a winged booster is tri-propellant---I guess what I am trying to do is gin up a system where an expendable upper stage really is super expendable (besides the payload). Folks in the 70's may have felt similarly.

The fly-back booster nursemaids the upper stage, so what de-orbits is little more than pot-metal...no SSME mass power heads coming in at Mach 2 like Columbia's break up.

Don't know about Starship engines.

Seeing the Merlin upper stage nozzles expended irks me.

Unless you can figure out composite rocket power heads, I don't think you can avoid that. I mean, whatever you use needs to be strong enough to survive hyrdolox combustion temps.

 

[TheKutKu]

Any possibility of solid expendable upper stages? That's probably a cheaper, and easier to produce, option than a large, S-IVB-sized hydrolox one. Considering american RP1/LOX engines of the 70s, it's also not much less efficient than the US-state of the art of Kerolox engines (for vacuum, 295-300s isp, vs 300-310s at the time).

They also can be much denser than cryogenic/semi-cryogenic stages, potentially making side-mount easier or even maybe putting them and the payload in a reusable fairing/cargo bay embedded in the booster (like STS's PAM/IUS or Neutron).

An optional third stage would already be all but required even for a hydrogen expendable second stage for GSO insertions, so an expendable upper composite made of a solid stage+third stage may be an interesting alternative.

A smaller kick-stage could also do complicated orbital insertions for low earth orbit (well, solid stages can also do complicated maneuvers if you're leaning into the ICBM-state of the art, but that defeats the purpose of cheapness).

It's really not much different from how Ariane had a Hypergolic+small hydrogen stage instead of Europa III's large hydrogen stage (sorry for the reference, being european it's the first thing that comes to mind)

Of course, due to the lower ISP (compared to hydrolox) or mass fraction (compared to kerolox), it's more advantageous for faster-staging boosters than slower ones, although the SRB's higher thrust will significantly lower the gravity losses compared to hydrogen stages; Since the first stage is reusable, it may also be worth it to moderately enlarge it to be able to carry a heavier upper stage to the same separation speed if it means the upper stage can be cheaper.

 

[Byeman]

Trying to split hairs there - transferring liquid propellant (fuel, oxidizer, or both) from one system element to another *IS* crossfeeding, independent of whether both elements have engines or not. Show me a definition that supports your narrow view.

Crossfeeding for launch vehicles is burning propellant from another stage/booster separate from the stage that the engine is mounted to. And continue to burn while the propellant supply is switched between them. Shuttle/ET or Atlas sustainer/boost package connections are not cross feeding. There is only one source of propellant and at no time is it switch. Only the propellant flow is stopped and then the connections are broken.

 

[martinbayer]

Crossfeeding for launch vehicles is burning propellant from another stage/booster separate from the stage that the engine is mounted to. And continue to burn while the propellant supply is switched between them. Shuttle/ET or Atlas sustainer/boost package connections are not cross feeding. There is only one source of propellant and at no time is it switch. Only the propellant flow is stopped and then the connections are broken.

Once again, show me an independent definition from a reputable source that supports your narrow view.

 

[Byeman]

That's exactly where I'm coming from - the ET and Orbiter simply were two separate and different system elements!!!

And I'd argue that feeding LOx and LH2 from the external tank to the engines on the orbiter is definitely crossfeeding.

Nope. One propellant source to one set of engines. Cross feeding is supplying propellant/fuel from multiple tanks.
The OMS engines could burn propellant from tanks in either pod.

Aircraft Fuel Systems | SKYbrary Aviation Safety

An aircraft fuel system enables fuel to be loaded, stored, managed and delivered to the aircraft engines. Its size and complexity may vary greatly depending on the aircraft type.

skybrary.aero

The "cross" in cross feeding is from fuel lines "crossing" each other when coming from different tanks.
It is not from "crossing" outer mold lines.
An aircraft drop tank is not cross feeding if it is only going to one internal tank

Shuttle MPS is not cross fed, it is fed directly from and only the ET

 

[Byeman]

Trying to split hairs there - transferring liquid propellant (fuel, oxidizer, or both) from one system element to another *IS* crossfeeding

not applicable to the shuttle. it is not "transfer". Propellant directly flows into the engines. There are no intermediary tanks.

Also, not "from one system element to another". It is one system, the Space Shuttle MPS.

 

[Byeman]

Now if an upper stage is methalox but small, and a winged booster is tri-propellant---I guess what I am trying to do is gin up a system where an expendable upper stage really is super expendable (besides the payload). Folks in the 70's may have felt similarly.

The fly-back booster nursemaids the upper stage, so what de-orbits is little more than pot-metal...no SSME mass power heads coming in at Mach 2 like Columbia's break up.

Don't know about Starship engines.

Seeing the Merlin upper stage nozzles expended irks me.

no, unworkable. The flyback booster would have too much energy to fly back

The Merlin upper stage nozzle is less material than the Merlin booster engine nozzle

 

[GeorgeA]

Based on the 1971 MDAC Expendable Second Stage study, from back when the Shuttle Booster was going to be a flyback stage and the Orbiter carried its own tanks, and a 65klb payload, you're looking at something like three-eighths the total vehicle mass

• Booster - new LOX/LH2 engine(s), total 1,770 lb thrust, 960klb propellant
• Upper stage - 2 x J-2 or equivalent, 290 klb propellant

That's a decent sized rocket, but not at all unreasonably so. The booster probably gets 3 x OTL SSME, or 4 smaller ones. Playing around with sizes, something like a 320klb engine would work with six on the first stage and one vacuum-optimised engine on the upper stage.

One F-1 derived engine would also work for the booster, but I think that LOX/LH2 would be preferred for 'efficiency', as it was with the OTL Shuttle.

If/when a small orbiter for this system was funded, it would carry maybe 10klb to orbit with 2-3 flight crew. Really a space station ferry craft; I'm assuming you'd get some kind of crew module for the payload bay.

The study I'm referencing assumed that the booster would recover downrange - Seymour-Johnson AFB in North Carolina was mentioned. For a low inclination launch I'm guessing you'd want somewhere in the Bahamas.

This sounds like it would be a useful system that would have had a long and effective life. Relatively straightforward to develop, simpler to transport and handle on the ground, and in accordance with the original Byeman requirement, well-suited for many of the missions of the 1970s-1990s that actually came to pass.

For propulsion, the "NASA 350K" variant of the XLR-129 might hit the sweet spot here. That would have also saved a few years and couple billion dollars versus (badly) reinventing the wheel at Rocketdyne. The upper stage application also parallels the Rockwell ESS S-II/SSME proposal.

Going back to the OP, I'm disappointed that the XS-1 never came to be. It seemed like a logical approach to operationally responsive small payload launches. I'm sure once SpaceX proved out the Falcon recovery and reuse process, it was clear there was no economical way to compete with them.

 

[GeorgeA]

Reviewing the ESS studies (as one does), I was struck by a few things:

• A fully fueled B-9U booster with the 550K SL/620K vacuum SSME version and a full-length S-II derived ESS with two of the 620K engines could have placed 250,000 to 300,000 lbs. in orbit. The largest payload NASA was interested in within the bounds of the study was the 176,000 lb. MDAC 33-ft-diameter space station module. For these lower-capacity launches, the ESS would have had reduced LH2 tankage compared to the stock S-II and the booster would not have been fully fueled for this type of mission.
• One of the more difficult problems was determining how far to offset the booster and the ESS when stacked. Too high and the center of gravity shifts would have needed to be compensated for by some serious gimbaling of the booster engines. A closer offset, however, would have led to significant acoustic issues for the ESS during the boost phase. Either way, it seems like the ride up would have been squirrelly AF for the booster pilots.
• I thought I knew a lot about this program, but I came across a truly whackadoodle idea: after orbital insertion, a booster/orbiter mission would have been launched to detach and recover the SSMEs from the ESS. As the orbital duration of the ESS was 24 hours, this had to happen within hours of the launch.

What is clear is that the study was comprehensive and, especially for our proposed smaller vehicle like the Yellow Palace booster, showed the concept was entirely feasible.

 

[Yellow Palace]

One of the more difficult problems was determining how far to offset the booster and the ESS when stacked. Too high and the center of gravity shifts would have needed to be compensated for by some serious gimbaling of the booster engines. A closer offset, however, would have led to significant acoustic issues for the ESS during the boost phase. Either way, it seems like the ride up would have been squirrelly AF for the booster pilots.

I suspect the vehicle configuration would be slightly different if the baseline was booster + expendable stage, so this might be mitigated for the notional 65k vehicle we're considering, at the expense of making it less optimal for a hypothetical future crewed version.

I'm definitely leaning to 6+1 new LOx/LH2 engines in the 300klb class for this vehicle. A pair of F404s might work for recovery propulsion on the booster; IIRC the B-9U had four F100s. Scaling up wouldn't be easy, but AFAIK it was a long time (if ever) before the 65klb payload became inadequate.

 

[publiusr]

The booster would fly-back. Upper stage expendable.

As for a hydrolox upper stage--has any thoughts been given to a coating that comes off with the fairings, so you could use the upper core as a telescope if not a whole wet workshop?

No foam to popcorn--propellant now leaving so less of a burst threat... maybe beam energy to the upper stage now that it is bare.

 

[Scott Kenny]

The booster would fly-back. Upper stage expendable.

As for a hydrolox upper stage--has any thoughts been given to a coating that comes off with the fairings, so you could use the upper core as a telescope if not a whole wet workshop?

No foam to popcorn--propellant now leaving so less of a burst threat... maybe beam energy to the upper stage now that it is bare.

No, hydrogen is a deep cryogen, you gotta have some insulation on the fuel tanks.

 

[Byeman]

The booster would fly-back. Upper stage expendable.

As for a hydrolox upper stage--has any thoughts been given to a coating that comes off with the fairings, so you could use the upper core as a telescope if not a whole wet workshop?

No foam to popcorn--propellant now leaving so less of a burst threat... maybe beam energy to the upper stage now that it is bare.

no. The insulation needs to stay on for the duration of the upperstage mission.
Beam what energy?

Wet workshop does not work. It makes the vehicle unworkable. It uses up payload mass in scarring the tanks. Do you not understand:

https://wikimedia.org/api/rest_v1/media/math/render/svg/f4f0e5fa252e0c5f23c387f1bfdf6784017907dc

You either have a launch vehicle or station module. Combining them compromises either task and increases final mass (m sub f). Which means to keep the same delta v, payload mass has to be reduced. "Wet workshops" will not be viable until there is infrastructure in orbit that can take stock stages and utilize them, without affecting their prime purpose as launch vehicles. Just as things are repurposed on earth, they are not scarred for the repurposing.

 

[GeorgeA]

I suspect the vehicle configuration would be slightly different if the baseline was booster + expendable stage, so this might be mitigated for the notional 65k vehicle we're considering, at the expense of making it less optimal for a hypothetical future crewed version.

I'm definitely leaning to 6+1 new LOx/LH2 engines in the 300klb class for this vehicle. A pair of F404s might work for recovery propulsion on the booster; IIRC the B-9U had four F100s. Scaling up wouldn't be easy, but AFAIK it was a long time (if ever) before the 65klb payload became inadequate.

P&W was far along with turbopump testing for both the 250K and 350K versions of the RL20/XLR-129 by this time. Again, if Rocketdyne isn't gifted the SSME contract and the requirement is a 300K reusable engine, SSME development should go much more smoothly and PFRT should happen by 1975.

 

[Archibald]

And, big irony related to Ares 1, XLR-129 was air-startable... unlike SSME. Because ISINGLASS RHEINBERRY origins.

 

[TheKutKu]

What would be the production rate/cost of a developped XLR-129 vs J2S?

Both have comparable thrust and mass, XLR has somewhat better vacuum isp at similar expansion ratio (~10 ish seconds), but wouldn't it be more expensive?