Could validation of transpiration cooling enable reuse of sustainer core rockets

[Stan-1967]

Summary of Premise:
I debated where to put this thread.  I didn’t want it to go into SpaceX, although SpaceX is clearly relevant.  It is more about re-use vs. any specific commercial or government launcher.  I think the principle of the topic could be relevant to may different launchers, IF it ever actually becomes viable.   To illustrate the topic with real world examples, I am thinking of mainly Ariane 6, but also SLS.   It could also apply to any other launcher using a center sustainer core with SRB’s for boost. ( H3, Vulcan, GSLV)

The premise boils down to this:  If transpiration cooling gets accepted & validated to be a viable technology for protecting an entire launch vehicle during a hot re-entry, does that create a more attractive entry point for launch vehicle makers to implement re-use on existing sustainer core architectures?  I have seen a general sentiment here on NSF that sustainer core architectures are a dead end technology for rocket designs.   The only supposed alternate is trying to follow more closely in the footsteps of SpaceX or Blue Origin & be in 3rd or 4th place before even starting.  I am questioning if that is indeed the case, or have other options presented themselves?

 The established paradigm for booster reuse looks like low velocity staging of the booster core, followed by a large stage 2 that is either disposable ( Blue Origin NG & F9) or highly complex & reusable  (SH+Starship).  However the sustainer cores of vehicles like A5/A6, SLS, & Vulcan reach near orbital speeds, so the thermal challenges of returning that core would be more similar to what is being proposed for Starship.  If anything, the thermal challenge may be quite a bit more benign, as Starship is being designed for Mars & Lunar return velocities.   At depletion, a sustainer core is on something like an orbital or suborbital trajectory with a perigee that takes it back into the upper atmosphere, so velocity is around 7000 m/s maximum. (?)  Thus the suggestion that a transpiration cooling scheme is a pathway for re-use on these types of vehicles.

To use SRB’s or not to use SRB’s, that is a question:
I think the initial steps of a re-usable sustainer core would initially need to continue the use of SRB’s.   I know they are generally not liked for cost, safety, & re-use considerations.   However the rocket equation likes them, & I included them in this sustainer core concept.  The primary reason is that a sustainer core needs very high thrust side boosters.  All the operators of sustainer core vehicles use SRB’s because they need the high initial thrust while the core lightens to a manageable T/W by SRB jettison.    If they had access to low cost high power liquid engines like BE-4, Raptor, or even RD-180/191, they might choose to build single stick vehicles.  Even with such propulsion access, there is a limit to this, ( cost mainly)  Vulcan being a good example that is closer to a single stick/sustainer core hybrid. 

Using SRB’s also placates some key political constituencies, as well as eliminates the need to develop high power & technically challenging main propulsion engines like BE-4, or Raptor.  Engines producing around 1000kN will work just fine.  Going with SRB’s in the short term, while not ideal, does seem to enable re-use gains, and sets a lower price floor for the operational vehicle.

As an example, the A6’s P120 SRB’s offer a pretty significant cost reduction vs. the P238’s on the A5.  I see estimates of 7-8M Euro’s per P120 SRB.  This is around 33% the cost of the 6-4 configuration, & 20% the cost of the 6-2 configuration.  If the upper stage of A6 cost another $10-$15M Euro, per launch prices for A6-2 & A6-4 could drop by around 50%-65% of existing prices, assuming at least 10-20 reuses per core.    This may be a price that sovereign entities would be willing to pay for a very long time in the face of pressure from cheaper foreign or even domestic suppliers.  SRB’s also scale well for any government led high value, high frequency launch requirements.


Examples to contemplate:
Again, not to focus on any particular vehicle, but the Prometheus engine & Themis development vehicle proposed by Airbus Safran is a decent example.   Currently, per the Themis videos on the Airbus website, they are looking to just copy the SpaceX model of booster re-use with a roughly 1000kN methalox engine.  However think of an Ariane 6 sized sustainer core, using transpiration cooling, with 3 to 5 methalox Prometheus engines which are sized about right to land a center core that weighs around 70t.  The higher density methalox vs. hydrolox in the same or similar core volume as A6, with SRB’s could boost total lift to LEO around 30%, which suggests there is margin for re-use operations while preserving existing payload capability. 

Airbus could properly expand the Adeline re-use scheme, and fly the entire core, equipped with transpiration cooling re-entry technology, to a preferred landing site far downrange off the coast of Africa.  The basic concept of operations is that after main stage separation, the small upper stage is already nearly orbital, & it proceeds to inject the payload into the desired orbit, however the lightened up sustainer core now restarts its engines to execute a burn to the target re-entry point near the landing site.  This is more work & more technically challenging than copying SpaceX or Blue Origin’s paradigm scheme.   It does fit better within Airbus’s industrial base, as well as propulsion choices.  It may be good enough to give cost savings that keeps their business viable longer term, as well as be a meaningful demonstration of technology leadership.

I find it interesting to apply the same thought process to SLS, especially in light of the recent political push for the moon by 2024, and the very sharp barbs poked at the prime contractor, Boeing.  I see nothing in the POR architecture that makes me think that vehicle is going to be fly past the first 4 launches, and even though I expect it to work, it doesn’t have the flight rate or operating costs to scale into a larger campaign for Lunar or Mars missions.   Some other entity like a SpaceX, ULA, or even Rocketdyne could really rock that SLS boat with a new re-usable SLS core design that also integrates the SRB’s & ICPS.  None of them will for their own political or vested ownership reasons, but how would a re-imagined SLS architecture with a re-usable core look to congress if the real per launch price was dropped to under $200M with a scalable flight rate?  ( $100M for the SRM’s, $50M upper stage, $50M or less for the amortized core).  Would the US Congress want to fund missions for that vehicle vs. rely on emerging commercial entities they can’t control?  Maybe?

Summary & roadmap forward:
Launch vehicle operators using sustainer cores should focus on the more interesting technology, which is transpiration cooling of the primary core.  In cases that need to switch main core propulsion from hydrolox to methalox or kerolox, smaller engines producing around 1000kN are more ideal & affordable than ambitious high thrust FFSC or ORSC designs.  Again, Airbus’s “Prometheus” engine is a good example.  Keeping existing technology like SRB’s for now is a conservative approach that allows the scheme to start gradually by making an acceptable vehicle that can be iterated upon from lessons learned while actually flying.

If an entity can master re-entry of a sustainer core, they are well down the path of learning what is needed for full re-use, & multiple roadmaps exist to achieve that ends. 

1.    Develop more powerful main propulsion & move to large multi-engine core ( like NG or SH/SS)  that stages low & slow for S1, & S2 re-use is solved from your prior investment on the sustainer core model.  This path seems best for development of Heavy lift architectures.

2.    Continue with a sustainer core model & develop small re-usable LRB’s that jettison at anywhere from 90-120 seconds & replace the SRB’s.    A Kerolox side mounted LRB is a potential good fit, & may yet have technical relevance as either a flyback, parachute recovery, mid air capture, or some other yet to be invented scheme.

I appreciate any thoughts & criticisms.

[rakaydos]

Sustainer cores have other problems for reuse. As we've seen on the Falcon Heavy Center Core, which is almost a sustainer stage itself, boostbak is basically not an option, and the trip back to the launch site is risky.

[Stan-1967]

Yes,  there would be no boostback.  It would be much more of a boost forward to downrange ( & crossrange) landing.  Sustainer cores like A5/6 or SLS are much higher velocity’s than FH center core.  They are nearly, if not orbital.  Post core shutdown burn could span continents.  Depending on the vehicle,  maybe even land back at the launch site after 1 orbit.  This is not SSTO,  the core is like stage 1.5 from the SRB contribution.