Friends, Americans, Countrymen. Lend me your ears. I’ve come to provide constructive suggestions to SLS, not to bury it.
A couple days ago, a friend asked me what I knew about the DUUS “Dual-Use Upper Stage” for SLS, and if there was any super-sekrit info on NASA Spaceflight’s L2. I finally came up for air yesterday from some proposal writing, and I’m supposed to be working on my USML Category XV NPRM comments, so of course in a moment when I badly needed a distraction, I decided to take a quick look. While unless I missed something, L2 doesn’t seem to have much more info on DUUS yet than the public side. But what I saw got me thinking a bit.
What we seem to know about DUUS is that it’s an upper stage for SLS, that uses an 8.4m diameter LH2 tank, a smaller diameter LOX tank (5, 5.5, or 6.5m diameter), and 4x RL-10 LOX/LH2 rocket engines. At least from how I’ve seen it described/depicted, it kind of looks like a super-DCSS (Delta Cryogenic Second Stage–the stage that flies on the Delta-IV), with separate tanks and a truss structure tying them together (sort of like how Xoie was arranged). Personally, I’ve always been much more of a fan of the RL-10 than the J-2X, and putting some thought into a non-ridiculous upper stage for SLS is long past due.
As an aside, back when the Augustine Committee was going on, I decided to run some architecture analysis numbers. One of the key takeaways I had was that especially for a chemical propulsion architecture, what you really want is a big, high propellant mass fraction, high Isp upper stage, with the ability to top it up after the fact. How that upper stage gets to orbit, and how full it is on the launch pad or when it initially arrives in LEO are really secondary concerns. While an existing Centaur stage looks like it might be able to do some bare-bones deep-space exploration when combined with on-orbit refueling, a bigger stage that didn’t sacrifice the Centaur’s pmf (90%) or Isp (452s-462s depending on the model) would be very interesting. NASA has wanted to do a large upper stage for SLS, but the question has always been if they could do a large stage that was also efficient. Previous Cryogenic Propulsion Stage concepts we’ve seen batted around had pmf’s as low as 75%, and used J2-X rocket engines, so while they were bigger, they didn’t end up anywhere near as efficient of designs as the Centaur. When I saw that DUUS was going to be an RL-10 based stage, it piqued my curiosity a bit.
Getting back to the topic, and to my “random thought”, I saw that someone had asked the question on the public side, if any of the DUUS concepts used a common bulkhead, like the Centaur propellant tanks. Most of the reaction seemed to be that it was impractical to do a common bulkhead design when the two tank diameters were not the same. Now, I was pretty sure I had seen some Centaur variants (the ones for Shuttle and Titan-III/IV) that had larger diameter LH2 tanks than LOX tanks. Sure enough after digging I found some verification in this document on pages 2 and 3. So yes, you can definitely fly a common bulkhead upper stage where one of the two tanks is smaller than the other.
Here’s a picture of one of the Titan-IVB Centaur stages being prepared for flight:
The Titan-IVB Centaur had been derived from the Shuttle Centaur design (Centaur-G and Centaur-G’) which had been cancelled after Challenger. I had correctly remembered that the wider LH2 tank was in order to shorten the stack, since the shuttle payload bay was a lot wider than the 3m Centaur standard diameter. But I didn’t recall why they widened the LH2 tank and not the LOX tank. After all, as commenters on the NSF thread were pointing out, having both tanks the same diameter with a common bulkhead is a lot more structurally efficient. Also the larger aft bulkhead makes packaging engines and other hardware easier–Have you ever looked at the tail end of a flight rocket upper stage? Those things get busy fast!
The answer that was provided in the Titan-Centaur document I linked to earlier was what gave me my crazy idea (emphasis mine):
The Titan IV Centaur, shown in Figure F, consists of a 576 cubic foot aft liquid oxygen tank that shares a common intermediate bulkhead with the 1890 cubic foot forward liquid hydrogen fuel tank. To reduce overall length, the LH2 tank diameter was increased to 14 ft. compared to the 10 ft. diameter Atlas Centaur tank, taking advantage of the 16 ft. diameter Titan IV Payload Fairing (PLF). A conical transition section joins the LH2 tank with the 10 ft. diameter LO2 tank, which attaches to the Titan Forward Skirt Extension (FSE). This transition ring maintains the successful 10 ft diameter common bulkhead, which reduced development cost and risk.
If you look at where the development cost and risk are, the common bulkhead, and the arrangement of engines, RCS systems, pressurization systems, etc. on the aft bulkhead are far more expensive and risky than the rest of the tank walls themselves. This is why you notice that over time the most common Centaur modification is a barrel-stretch. Adding a little length to the barrel takes a little structural analysis, a bit more pressurant, and probably updating the GN&C based on new vibration modes and such. But going to a new diameter, or going to a new engine configuration is a lot more expensive and time-consuming, because there are a lot more items you need to requalify. Going to a two-engine Centaur design for Commercial Crew launches is only really affordable because Centaur used to fly with two engines, and going back to that design (while rolling in the upgrades they made for the Single-Engine Centaur) is easier to delta-qualify than a clean-sheet design.
So with all of that prologue, my crazy idea was what if you designed the DUUS as a common-bulkhead stage, where the LOX tank, common bulkhead, and aft bulkhead/engine bay area were the same as the 5m diameter LOX tank, common bulkhead, and aft bulkhead/engine-bay of the planned ULA ACES upper stage?
Why would you want to do that? Here’s some potential benefits:
- Even with the less structurally efficient taper-section, the design will likely be more efficient than a DCSS type separate bulkhead arrangement. Structural efficiency matters a ton for upper stages.
- A common-bulkhead design will also be shorter, enabling either more volume for payloads in the PLF, or less bending moments on the vehicle.
- Compared to an equal-diameters common bulkhead design, there will be significantly smaller whetted area between the LH2 and LOX tanks–this will reduce the heat leak from the LH2 to the LOX.
- It’s a lot easier with a common bulkhead design to use LH2 boiloff to reduce and/or suppress entirely the LOX boiloff. Hydrogen has a much higher heat of vaporization and heat capacity than LOX, and even boiloff hydrogen is much colder than the freezing point of LOX, meaning that you lose a lot less mass overall if you’re using the hydrogen as a heat spongue. And this is easier to do with a common bulkhead design.
- Using the “bottom half” of an ACES stage means you’ll also be getting the Integrated Vehicle Fluids system, which is a high-efficiency way of providing long-duration power, propulsive propellant settling, and repressurization.
- Not only is IVF more efficient, since it’s using LOX and LH2 from the main tanks for those three functions, you don’t need bigger pressurant bottles or RCS bottles for the two designs. You can either run the IVF system at a higher duty cycle, or maybe add extra IVF modules to deal with the vastly different tank volumes between an ACES stage and a DUUS.
- Manufacturing Commonality! If done right (no guarantees on that), the DUUS could use either an identical or a modularly upgraded ACES LOX tank and aft bulkhead design, which could be produced on the same line as the rest of the ACES stages. This should lower the marginal cost of the system by leveraging larger efficiencies of scale. The LH2 tank would be based on the Michoud SLS core stage tooling/fabrication equipment, meaning that even if you didn’t need a DUUS for a while, the capability wouldn’t just disappear. You don’t have anywhere near as much of a standing army that is DUUS-specific that has to be constantly fed even if you don’t need the DUUS at that given time.
- Design Commonality Using this approach means that you can get two great upper stages for about the cost of 1.25 stages worth of development. If NASA can take a light hand on the requirements side (stop laughing!), the DoD might be willing to split the development cost, because having the ACES stage could be very beneficial to the DoD.
- By helping get the ACES stage qualified, if you need a high efficiency third-stage on the design, for deep-space missions, you’ve got one developed for free, and being used commercially. And it will be much better than the DCSS-based iCPS that they’ll be using for the first SLS flights.
On the ACES side, getting NASA to cover some of the development costs means that ULA is more likely to be able to close the case for fielding ACES, which could provide some real benefits to ULA, DoD, and NASA:
- Having an ACES stage could enable larger payloads on Delta-IVH and could boost the payload capacity of Atlas-Vs without strapons that you might be able to eliminate the Atlas-V strapons entirely, saving some money.
- Or if you keep the Atlas-V solid strapons, the payload on an Atlas V 552 (though I guess it’d be a 554 in this case) would be high enough that you might be able to eliminate the entire Delta-IV line.
- Even without refueling, Atlas V Phase 1, which is what this would be, is a very capable vehicle for LEO, GEO, and deep-space missions. Not everything is big enough to justify SLS.
There are probably plenty of other benefits I’m not thinking of.
What about the challenges? Here are a few I can think of:
- ULA has moved back to the current CRES material used in Centaur for their ACES concept. The Michoud 8.4m diameter tank line uses Aluminum. Either 2219 or 2195 Lithium-Aluminum. This would require some sort of dissimilar metal connection where the taper section ties into the LOX tank. As I understand it, ULA has done some work already on explosive welding aluminum to their stainless material, as part of some of the common bulkhead work they’ve done in the past for ACES. While this is probably solvable, throwing some analysis and early prototype testing at this joint is probably a good idea.
- It doesn’t use the J-2X, which we’ve now spent a lot of time developing. The J-2X just isn’t a good upper-stage engine. Its Isp isn’t that spectacular (448s vs 451-462s for RL-10), and it is just way too big for most upper stage applications–a J-2X engine weighs more than the entire dry weight of a Centaur upper stage. You just don’t need a high stage T/W ratio for most upper stage applications. It might be ok for an explicit second stage that was only used to get something into LEO, but for going beyond you really want an optimized upper-stage system that’s as sleek and efficient as possible. Leaving J-2X on the shelf for other applications is better than compromising a design just to use a politically-favored component. Aren’t we doing that enough with SLS?
- Will NASA actually be able to take a light-enough hand on requirements to enable their DUUS LOX-tank and aft bulkhead/engine-bay to stay common-enough with the ACES hardware to get any net benefit from commonality? I hope so.
- Wouldn’t having both tanks the same diameter save even more weight and height, and provide more aft-bulkhead real-estate, enabling more engines if necessary, and generally just being better? Maybe. But at the cost of NASA having to qualify a bigger common-bulkhead, losing all commonality with commercial launchers, and having to foot the whole bill for the DUUS standing army. Is that worth a slight performance gain?
- Isn’t it impossible to fit 6 engines onto a 5m diameter stage? It probably depends on which RL-10 type you choose. The earlier ACES stages showed 6-engine options, but the more recent ones seem to focus on 4 engines, possibly with bigger nozzles. This is a good question though.
- Isn’t a taper from an 8.4m diameter tank down to a 5m tank a bit extreme? Glad you asked. I decided to make a spreadsheet to investigate this (had my VPN been working, I’d even have a SolidWorks 3D model tied to the spreadsheet, but for now you’ll have to live with the graphic in the third Excel sheet. Here’s what such a stage would look like with 300klb of propellant and a 30 degree taper angle:
And here’s what one would looke like with 450klb of propellant and the same 30 degree taper angle:
Neither of those looks particularly terrifying.
Anyhow, it’s a bit of a crazy idea, but I think it has some merit. Anyone who reads this blog regularly knows that there’s little love lost between me and the SLS, but this might be a way of doing things that if done well could actually be a real net-benefit to the community.