[Note: I wrote this article based on publicly available information, without consulting with the Quest guys, so any errors are probably my misinterpretations.]

But before we jump into the technology, let’s do a refresher on what MLI is for those not so familiar with the area:

What is MLI?

MLI on a Mars probe (Credit: NASA)

MLI or Multi-Layered Insulation is a form of thermal insulation that uses alternating thin layers of metalized plastic (typically Kapton or Mylar) and spacer nets (usually made of Polyester or fiberglass) to slow the radiative heat transfer into or out of a spacecraft in a vacuum environment.  When you see pictures of spacecraft covered in what looks like gold foil, that gold-colored stuff is usually MLI.  Wikipedia has more details here.  MLI has been around for almost half a century, and is one of the key elements of spacecraft thermal control.

MLI is one of the best thermal insulators known to man, but there are several hitches to existing MLI:

• The thermal insulation performance of MLI tends to be variable.  One of the ULA guys (Frank or Bernard) related an anecdote that they once flew two back-to-back Atlas or Titan missions, using MLI as the thermal insulation on the Centaur tank, and they actually had 30% difference in boiloff rates, in spite of solar levels and mission profiles being almost identical.
• MLI is very structurally fragile because it is basically a bunch of plastic thin films only held together at the edges.  This means that they can’t be exposed to flight aerodynamic loads without getting quickly destroyed.
• MLI only works in a reasonably good vacuum (under 1 mPa or 8×10^-6 Torr). When combined with the previous challenge, this means that even though they’re much better insulators than say SOFI (Spray-On Foam Insulation), they can’t be used for insulating external faces of launch vehicle propellant tanks.

Quest, working with Ball Aerospace, has come up with a clever technology which they call Integrated MLI (IMLI), and several derivative technologies including Load Responsive MLI (LR-MLI), MMOD-MLI, and Launch Vehicle MLI (LV-MLI), which solve these weaknesses of traditional MLI, enabling many neat new space technology applications.

Integrated MLI

The core innovation that Quest and Ball came up with was the idea of replacing the plastic netting “scrim” layer with evenly spaced “micro-molded” snap-together polymer supports.  These micro-molded supports keep the MLI layers consistently spaced, transfer loads, and greatly reduce the conductive heat transfer between MLI layers.  To give you an idea of what these things look like, here are some pictures (borrowed from this brochure made by Quest’s micro-molding partner, Phillips Plastic Corporation):

[Editors Note: I'll have a picture here later when I can fix an upload bug with WordPress.  For now you'll have to read the brochure linked above to see what I was going to put here]

The neat things I see about this approach are:

• The IMLI blankets are now a lot more thermally deterministic, repeatable, and analyzeable.
• The IMLI blankets have much smaller thermal conduction contact area between each sheet, demonstrating around 30% better insulation than traditional MLI of a similar number of layers.
• The micro-molded snap elements tie the layers together and are anchored all across the surface you’re trying to insulate, instead of just along the edges, making IMLI significantly more robust than traditional MLI.
• By replacing the scrim layer netting with a few discreet micro-molded pieces, they’ve probably cut weight compared to traditional MLI blankets of the same number of layers.
• There’s some real potential for mass production and automated assembly that could drive down costs significantly.

And the technology behind IMLI also serves as the foundation for the other three derivative technologies.

Load Responsive MLI

While Integrated MLI was a big improvement over traditional MLI, you still could only use it in vacuum environments.  Quest and Ball developed what they call Load Responsive MLI (LR-MLI) to enable customers to have the benefit of MLI even in an atmosphere.  Basically, LR-MLI consists of a thin vacuum shell supported by some spring-loaded spacers, with a vacuum pulled on the space between the vacuum shell and the underlying structure. When the external pressure is non-vacuum, the spacers are forced flat, where the center of them rests on the center of the spacer below them.  This increases the heat leak through the spacers, but still provides a much better insulation than SOFI (their .25in thick test part provided better insulation than a 16in thick layer of SOFI!).  Once the external pressure starts falling off, the spacers push back apart in a way that greatly reduces the conduction path, resulting in a really good thermal insulation on orbit.  See this page and page 22 of this FISO presentation for illustrations of the concept.

Benefits of this approach as I see it:

• You now have a non-SOFI method for insulating a tank that works in both atmosphere and in-space that doesn’t have the popcorning problems SOFI has, which both eliminates debris falling off during launch, and also eliminates the risk of insulation flaking off once in orbit.
• LR-MLI is a significantly better insulator both from a mass and a thickness standpoint compared to SOFI.
• You get rid of the need for GHe or GN2 purges on the ground.
• Enables fairly lightweight dewars to be constructed for applications that need it.

My only concern is the challenge of maintaining a vacuum for a long duration on the ground, though I guess dewars are used a ton in industry, so maybe this isn’t a huge deal.

Launch Vehicle MLI

The latest improvement on the IMLI theme, for which Quest just finished a Phase I SBIR contract for last year, is an MLI technology capable of being used on external aerosurfaces of launch vehicles.  This Launch Vehicle or LV-MLI appears to be a combination of the LR-MLI concept with a thin aeroshell surface.  There aren’t as many details on the concept, since it’s still in active development (here’s to hoping things went well and they get a Phase II award next month!), but here are the SBIR abstract and briefing chart.  The goal is to have an insulation system that weighs about a third of what a 1.9cm SOFI layer would, but with 85X the insulation value.

MMOD-MLI

One other related concept that Quest and Ball developed is an IMLI variant that includes integral MMOD (Micro-Meteor/Orbital Debris) protection capabilities.  This MMOD-MLI includes layers of Kevlar and Nextel cloth between layers of insulation, providing the same sort of multi-shock shielding capability that is what makes Bigelow’s modules so much more robust than older ISS designs, while still packaging things in a neat, multi-functional structure.  Basically, an incoming piece of MMOD would hit the outer layer, instantly vaporizing the MMOD, which would then have its energy absorbed and the momentum distributed as it passes through the multiple shield layers.  Like LV-MLI, MMOD-MLI just finished a Phase I SBIR a few months ago, so the results aren’t all out, but the goal was a design that would give a propellant depot a 95% chance of surviving its design lifetime without an MMOD-induced failure, without adding substantially to the MLI mass, or significantly decreasing the MLI thermal efficiency.  The neat thing about this technology is that it looks like it can be integrated with LV-MLI or LR-MLI without much additional effort.

So you could theoretically get an MLI shield that can function in both atmospheric pressure and on-orbit, could take aerodynamic loads, and when on orbit could double as a very effective MMOD shield.  Think about that one.

Some Random Applications

This is far from all the space applications enabled by these technologies, but here are a few less-obvious ones that I think are worth mentioning:

1. Cryogenic-fueled Air-Launched Rockets: One of the big challenges for externally-carried cryo-fueled air-launched vehicles (including even LOX/RP-1 designs) is that the heat transfer environment during the flight from the ground to the launch point is substantially worse than a vehicle experiences on the ground, due to convective heat transfer from air flowing over the launch vehicle during flight, which may very well be an order of magnitude or more than what is experienced on the ground.  And unfortunately, air-launch vehicles are typically much more sensitive to losses due to boiled-off propellants. The traditional thought on how to handle this is to have some sort of Airborne Service Equipment (tanks and plumbing and stuff) that either keeps the tanks topped up, allows you to only load the cryo propellants at the last second, provides some sort of sacrificial coolant, or provides an active cooling loop. With something like LV-MLI (or LR-MLI inside a separate aeroshell if your tanks aren’t conformal with the outside of the vehicle), you could cut down on the heat leak substantially.  Maybe to the point that you could eliminate or greatly simplify the required ASE complexity, cost, and weight.  Maybe combine pre-subcooling the propellants a bit with the insulation and you might be able to get rid of the ASE requirements entirely.
2. Wet Stations: One of the ideas that made the rounds a lot a few decades ago was that NASA should haul the Shuttle External Tanks all the way into orbit (instead of ditching them at just below orbital velocity to burn up in the Indian Ocean).  Two of the technical issues with this idea were that the insulation on the ET was liable to flake and pop off in orbit, potentially creating space junk issues.  There were probably solutions to this problem, but they likely involved either a lot more mass, or a lot of added complexity compared to just using an insulation system that isn’t prone to flaking off.  The other issue is that I don’t think SOFI makes a very good MMOD shield, meaning that a structure that big had a pretty likely chance to have an MMOD failure during its lifetime if some external MMOD shield wasn’t added.  Using a combined LV-MLI and MMOD-MLI solution, you could lower the weight of the insulation system overall, increase payload on normal flights, and completely eliminate this problem.  And this isn’t just limited to SLS, this could also be the case for any other core stage or large upper stage that reaches near orbital or orbital speeds, such as the core stage on a Delta-IVH or eventually an ACES upper stage. As an added bonus, you could even get the 60s-Retro Black-and-White stage coloring scheme without the weight penalty.
3. Bigger Single-Launch Propellant Depots: Along a similar vein, this approach could allow you to do one of the ULA single-launch, dual-fluid depot concepts where the LH2 tank is built into the upper stage’s payload fairing outer mold-line, enabling a 70-75mT LOX/LH2 capacity using Atlas’s Centaur, or over 100mT using Delta-IV DCSS as the starting point.  The nice thing is that not only do you get lightweight, high quality insulation, but you also get MMOD protection at the same time (which is critical for a depot).
4. Super Jumbo Single-EELV-Launch Propellant Depots: On the crazier side, you could combine ideas number 2 and 3, and say have a Delta-IVH place its core stage (with an LV/MMOD-MLI combo in place of its current SOFI) into orbit, with what payload remains being a docking node, temporary stay habitat, or additional propellant tanks for other more storable propellants if you want a multi-propellant depot.  That gets you up over 200mT of LOX/LH2 capacity in a single launch, without requiring an HLV to do it…  Though admittedly, a Delta-IV CBC would take a lot of modifications to get the passive cooling right compared to the Centaur-derived approach that ULA proposed originally.

First off, I wanted to congratulate my friends at XCOR and ULA. This engine work that ULA and XCOR have been doing is something I’ve been watching from the sidelines for some time now, and it’s cool to see them making progress. As Jeff Greason pointed out during and after the Augustine Committee’s work, the US rocket industrial base is in bad shape, and getting new blood and new ideas injected into it is critical.

Second off, I’ve been an advocate of aluminum rocket engine fabrication for several years now. It’s worth noting that while I was still at Masten we ended up doing almost every one of our successful Xombie/Xoie flights using aluminum chambered engines (and I think we’re still the only company to ever fly a reusable rocket engine made of aluminum). Aluminum has a ton of advantages especially for cryogenic engines (ie Methane or LH2 fueled ones), but even for non-cryo ones as well. A quick list includes:

• Low density and high strength-to-weight allows you to get a very lightweight engine without having to push margins or analysis anywhere near as far as with more traditional materials.
• Low-cost and easy availability of many alloys with good mechanical and thermal properties. Once you’ve tried to source a high-strength copper alloy for a medium-ish sized rocket engine you’ll know why this matters.
• Easy, quick, and cheap to machine, even if you want to do tricksy things with the cooling groove geometries.
• There are a ton of manufacturing process options that are semi-unique to aluminum that give you a lot of tools for optimization of the design without excessive costs. Some of these knobs allow you to optimize either for maximum heat flux into the coolant (for expander cycle engines) or minimum heat flux into the coolant while still keeping the wall cool.
• High thermoconductivity (about 50-60% of pure copper’s thermal conductivity) allows you to keep walls cooler–which is kind of necessary with it’s low softening temperature.
• If you can keep it cool enough for long-duration operations (which you usually can for low-moderate pressure engines), thermal stresses can be much lower making it easier to make engines that can stand hundreds or even thousands of cycles

The list definitely goes on from there (like making feasible an alternative engine cycle that I was supposed to have blogged about months ago), but that gives you an idea. The manufacturability/availability issues were enough to get me an opportunity to try them out at Masten, and the work we did for the Xombie/Xoie engines vindicated the choice. For an upper stage engine, the benefits are even more compelling. One of the things I’ve always looked for are manufacturing technologies/choices that allow you to cheat on the cost vs. performance curve. With a small alt.space company, you’re not going to be able to spend the same amount of engine optimization as a bigger aerospace company, so any technologies that allow you to approach “big boy” performance while still being something that a 1-3 person propulsion team can do is worth pursuing.

I think this technology is especially relevant to RL-10 follow-on type efforts like what ULA and XCOR mention they are collaborating on in this announcement. Using the right combinations of manufacturing processes (and there are probably several ways of skinning the cat), you can increase heat flux into the coolant (which allows you to get more power out of the engine or higher chamber pressure), lower the weight of the engine assembly, substantially reduce the manufacturing/inspection/rework cost and complexity compared to a tube-wall nozzle, improve the reusability of the engine, and at the same time allow robust enough margins that a small team can have a realistic shot of delivering a world-class engine.

While I am very happy for XCOR and ULA, I do have to admit to being somewhat jealous that I haven’t had a chance to be involved in this aluminum nozzle technology effort. I spent a lot of time at Masten working on coming up with approaches for making scalable, low-cost, high-performance manufacturing approaches for aluminum nozzles, with just this sort of application in mind, but we were never able to get the sort of outside traction ($$$) it would take to actually validate our concepts (past what we did for the Xombie/Xoie/Xaero/Xogdor engines). Since leaving and starting Altius I’ve been trying to push the ideas even further. In fact, this past month I came up with a completely new approach that if it works (I’d give it about a 75-80% chance of working) could be amazing, not only for rocket engines but also for 3D printing, and many other applications as well. Imagine a process that would make a full-density part with lithium-aluminum strengths, where minimum hole size for internal channels was small enough that you could basically make metal foams, that would allow you to build-in electronic components and sensors, but without the size limitations of most other additive manufacturing processes, which could be scaled up for large thin structures (on the scale of an F-1 rocket engine or an Apollo CSM-sized transpiration-cooled heat-shield).

Anyhow, I hope that some day we’ll get to see some more details on what exactly XCOR/ULA doing for the manufacturing process, and I also hope that we’ll see an RL10-class engine flying some day with an aluminum nozzle (and maybe even chamber). Congrats guys!

“While it is true that prudent investments in science and technology will almost certainly yield future economic gains and will allow our knowledge economy to grow, it is also true that these gains can be thwarted by poor decision-making,” Chairman Hall said. “Americans expect and deserve better. With our unemployment hovering at over 9 percent, they expect us to reduce or eliminate those programs that are duplicative and wasteful and examine ways to advance real job creation and economic growth, not just spend their hard-earned money on what the government assumes is best for them.”

Hmmm…I can think of a few examples of massive projects that fit those descriptions. Like say SLS?

• Duplicative: In the near term, without any plan for BEO exploration hardware, SLS will be duplicating the function of commercial crew launchers–sending astronauts to the station (on MPCV), but at far higher costs.  If at some point an upper stage and actual exploration hardware do get funded (10-15 years from now), SLS will mostly be launching propellant–something private rockets are also capable of doing.  So, duplicative? Check.
• Wasteful: The development cost to the government of both the EELV programs and the two COTS programs are less than half one year’s worth of SLS funding, even at the lower projected rate that Obama proposed in his FY’12 budget, and even including the $300M increase in COTS funding. Those programs are giving NASA and the DoD four launchers, with at least some capabilities on both coasts, as well as two ISS cargo vehicles.  Even if you include the total government and private investment in developing those rockets, the total cost is far less than SLS alone will spend over the next 3 years, let alone to completion. Even if you insist on building an HLV, and even if NASA is the only customer (likely), these programs demonstrate that the expected $10B+ development cost for SLS is disturbingly high compared to the cost of developing vehicles for launching critical national security assets.  Wasteful? Check.
• Job Creation? Economic Growth? SLS is a NASA-specific products with no real outside commercial benefit, that will reuse obsolete technology in an effort to maintain as many existing jobs as possible. It is really just a zero-sum wealth transfer from the productive part of economy to politically-connected contractors. Actual economic growth and job creation come from creating new goods and services that provide for peoples wants and needs at increasingly affordable prices.  Innovation, both technical and entrepreneurial are what drive job creation and economic growth–not running government-directed design bureaus to produce products that are irrelevant outside of NASA’s needs.  Contrast this with EELVs, the COTS vehicles, Commercial Crew, and many of the technology programs NASA wants to fund, that serve multiple public and private needs, create wealth, and provide jobs that are backed by eventually self-perpetuating wealth-creating enterprises. So, Job Creation, Economic Growth? Not so much.
• Spending Tax Money on What Government Thinks is Best for Them? The big push for SLS development from the Senate (which has won it the nickname “the Senator Launch System”) and the even bigger push from the House, were led primarily by Congressmen from states that directly benefit from continued spending in this area. The ironic thing is that it’s pretty clear that even NASA doesn’t entirely want SLS, but you have Congressmen trying to legislate the design of a launch vehicle. To the point of Utah reps bragging about how language they put into the bill supposedly can only be met by using hardware procured in a non-competitive manner from bloated contractors in their districts (which I’m sure didn’t make any campaign contributions to help grease the palms of their Congressional enablers). There are few clearer examples of Congress forcing the government to build something that is more in the personal interests of certain Congressmembers than is actually beneficial for tax-paying public. Check.

Not trying to pick on Representative Hall.  I actually have a bit of a soft-spot for the guy, since he came out and spoke at the NGLLC awards ceremony.  Just pointing out that he has good advice, and it would be great if he consistently followed it.

Personally, I think this is mostly a terrible idea. While offering fixed-price services, and moving to FAA regs is nice, I really don’t see how this fits with the spirit of CCDEV.  After all, USA is talking about taking over an existing government asset, and flying it temporarily through 2017, not providing a long-term commercial crew capability for ISS in the post 2016 timeframe.  And the budget ($9B over the next six years), is way outside the $6B NASA was going to give to commercial crew, or the $3B that the anonymous Senate staffer last week thought would be the real number.

But shuttle huggers, don’t despair.  If something like this goes forward, they could probably do this by taking money from the SLS and MPCV budgets.  After all, this would be offsetting some of the carrying costs that NASA would have to pay for keeping the Shuttle infrastructure in place.  By doing this, there also wouldn’t be any rush to finish Orion or the 70-100 ton version of SLS, because you could just keep flying the shuttle “commercially” for another year or two if commercial crew faces delays.  In fact, this would allow NASA to go straight for their beloved 130mT SLS and deep-space rated MPCV, because there would be no need for the intermediate vehicle.  They can take as much time as they want.

The only even remotely legitimate purpose for trying to rush SLS/MPCV was the worry that possibly all of the commercial crew providers would be running late.  It’s possible I guess, especially if they try and put all their money on just one or two providers.  But, under the current Senate-designed plan, if commercial crew does work, SLS/MPCV would be a giant budget-sucking white elephant for several years while actual mission hardware (EDS stages, landers, and/or habs) was developed.

But with this plan, you can just go straight to “exploration class” HLVs and mission hardware, without having to worry about the fate of ISS.  Something like this would allow you to keep your HLV infrastructure alive until you actually need an HLV without killing commercial crew.

And anyway, SLS and MPCV have big enough budgets that this would only be cutting out maybe 1/3 of the money they’d be getting over that time frame.  If the DIRECT fanboys are right, there may even be a straightforward way for NASA to still deliver on something like that within the budgets they’ve been given, even with keeping the Shuttles flying.

And if there are budget cuts, hey you have the shuttle still flying, you can just stretch out the SLS development even further.

If NASA tries to go this route, they should do so under the SLS budget, not the Commercial Crew one.

Here is the humor slide I started out with:

We Don’t Need No Steeking Propellant Depots!

My actual presentation:

RLV Friendly Depots

Bernard Kutter’s presentation for ULA:

Near Term Depots

and Dallas Bienhoff’s presentation for Boeing:

Space Transportation Impedance Matching

I haven’t been given a copy of Rand’s presentation yet.

[Edit, here it is, Rand says he'll probably get some annotations up later]

Anyhow, a few quick random thoughts that I don’t think anyone else has really hit upon on the intarwebs:

• One of the concepts out of Dallas’s presentation I liked was the idea of having a space transfer tug that takes landers from EML1 (L2 would also work) to some perilune trajectory, and then returns to EML1.  I’ve been toying with variants of this idea for some time.  With a Centaur-sized trasnsfer tug, fully-tanked-up in EML-1/2, you can actually bring pretty darned big landers most of the way to the lunar surface (ie leaving 1000m/s or less of delta-V for the descent), while still having enough propellant to return to lunar orbit and from there to the L-point station.  That segment is probably one of the easiest in-space segments to start doing reusable stages, since you don’t need an aerobrake, and don’t have to deal with lunar dust, just propellant transfer, and lots of engine relights.
• In a conversation with Jeff Greason late one night at the conference, we got off onto the topic of RLVs and propellant depots.  One of Jeff’s opinions is that in order to really have an industry for some service, you need enough demand to allow for 2-3 healthy competitors.  With only one provider, you get monopolies, three is ideal.  But for RLVs you probably want a small fleet (~3 vehicles) of RLVs so that you can provide dependable service even if you either have a mishap or have to pull one of the vehicles for maintenance or repair.  Having a single vehicle may work during the development phase where you’re transitioning into operations, but once you’re in full operations, you want enough demand for 2-3 companies with probably 2-3 vehicles per year.  And for each of those vehicles, in order to get the per flight price in a really good range, you need to fly often–Jeff says 100 times per year, but I’ve heard numbers as low as 30-50 (but any way you slice it, it’s a lot of flights).  That comes out to somewhere in the 120-900 flights per year range.  The interesting thing that Jeff mentioned was that if you postulated very small RLVs to start with (say 300-500lb to LEO net payload capacity), just one lunar mission per year would be enough to provide enough demand for an entire healthy industry by itself.  Towards the lower ends of that scale, you’d only need one “soyuz around the moon” flight, or 1-3 GEO flights that used a propellant tank-up in LEO (say using a Falcon 1 with a mini-Raptor type LOX/LH2 upper stage?) to provide enough demand for at least the starting of an industry.
  • While 300-500lb to orbit sounds tiny, that’s actually a pretty reasonable size for a first-generation RLV.  The first stage doesn’t end up being that much bigger than existing or planned suborbital vehicles, doesn’t have to have much more capability either.  The upper stage ends up down in the middle of the size range for proposed suborbital vehicles.  While it has a much higher performance requirement, and much nastier reentry environment, it’s on a size that you can realistically work with a lot easier.  Also, a lot of the TPS work can be refined by flying “expendable” upper stages on these first generation commercial suborbital launchers.
  • This would definitely require the sort of RLV-friendly depot setup I described in my presentations–you’d have to have tugs that carry all the rendezvous/docking smarts, and keep the RLV-side of the propellant system as dumb as possible
  • Propellants are a much less demanding payload than people.  Not only does this keep up-front development costs down, but it also reduces the business risk if you happen to lose a vehicle occasionally.  While high flight-rate RLVs should be capable of high reliability, we’re also talking about 1st or 2nd generation systems here, where we’re still learning a lot–and learning can be painful.
• I also liked Bernard Kutter’s graphic of the simple, single-launch, dual-fluid depot concept.  This is a simpler version of the ideas Frank Z and I came up with last year (it uses a stock Centaur-sized tank for the LH2 side of the depot), but is still quite capable–on the order of 30mT capacity is nothing to sneeze at.  With one of those in LEO and one in L2, that’s actually enough to do an ESAS-capacity lunar transportation system without Heavy Lift.
  • One of the really interesting possibilities is that if something like this demonstrator depot were chosen as a part of the money Obama has proposed for orbital refueling technology demonstration (it wouldn’t need anywhere near the full $400M-1B that Obama mentioned per technology area), if the demo system worked, it would actually be operationally useful.  Sure, you’d want to replace it and/or upgrade it down the road with lessons learned, but I’m a fan of pressing technology demos into operational service, as that’s a good way to get a lot more data out of the deal.
• I also liked how Bernard explained a lot of the cryo storage issues.  A lot of this stuff still needs to be proven in space, but they (ULA, LM, and Boeing) have a lot more experience doing related tasks than most people realize.

I probably have some more thoughts on the matter, but I’m home at sick with a cold today, so I’ll leave it at that.

I was going to write a blog post on this earlier, but Jeff Greason beat me to the punch in comments over on SpacePolitics.com (emphasis mine):

Both the Phase I and Phase 2 versions can support 7.5m fairings; I’ve discussed the fairing size argument elsewhere and won’t repeat it other than to say that seems sufficient to me for a long time to come. To some that seems larger than needed, others envision a piece or two of hardware (such as the Mars manned entry vehicle) which we might need in the far future and which, assuming NO new technologies are applied between now and then, might require a larger fairing. Personally, I would rather start exploring soon and assess our needs again when we are a little closer to the point of need than contort our launch infrastructure today around this hypothesized future need.

…

Of course once we are using kerosene boosters and have kerosene engines, developing a larger booster, once there is need, presents no special technical difficulty. This was called “Phase III” in the EELV briefing to the committee in public hearings. Such a booster will cost more money (as all ultra-large boosters do), and there is limited forseeable market for it (as for all ultra-large boosters), but if it turns out we need it in the future, and are willing to spend money on it in the future, we can do so. There is no need to spend today’s scarce dollars on an ultra-large booster for which we have no near-term need, just in case we will need it someday. Those who have claimed this is an all-or-nothing decision are ignoring alternatives.

More thoughts later.

RLV Thermal Protection Systems
Protection from the harsh heating environment caused by atmospheric reentry is one of the biggest challenges for reusable vehicles–far more difficult than the often harped-on rocket equation or the “inefficiency of chemical propulsion”. The problem isn’t even the weight of the thermal protection system as much as it is the maintenance requirements. Ideally you’d like a TPS solution that requires very little maintenance, and can be “tested” easily and quickly on the ground before flight, even if it cost you a little extra weight. You’d also prefer something that was relatively simple operationally to use, with a minimum number of failure modes. MHD thermal protection seems like an interesting match for these requirements. I should note however that there are other promising ideas out there such as transpiration cooling that might also work on their own or in combination with MHD thermal protection, but they are beyond the scope of this blog post.

Some Take-Aways from the Literature on MHD Reentry TPS
There have been several interesting papers on this topic, including the JS&R article “Experiment on Drag Enhancement for a Blunt Body with Electrodynamic Heat Shield” that got me thinking about this more seriously, a second JS&R article that goes into experimental proof of the heat flux reduction “Experimental Verification of Heat-Flux Mitigation by Electromagnetic Fields in Partially-Ionized-Argon Flows”, and another JS&R article from a year and a half ago “Numerical Analysis of Reentry Trajectory Coupled with Magnetohydrodynamics Flow Control” that I’ll be leaning on pretty heavily for this discussion. You can purchase the articles from AIAA, or if you already have a subscription to the Journal of Spacecraft and Rockets, you can read them for free.

I’ll briefly summarize some of my takeaways before going into my thoughts on how to move things forward from there:

1. Both analytically and experimentally, magnetic reentry TPS appears to provide large reductions in both peak heating and in total heat load.  The third paper above suggested a 30% reduction in peak heat load and a 40% reduction in total heat load for ballistic reentries.  Under the conditions tested in the second paper, heat reductions up to 85% were shown.
2. The magnetic braking effects dominate aerodynamic braking effects at high altitudes.  This is mostly due to lower density meaning that atmospheric drag is fairly low, while also lower density means that Joule heating caused by the currents (the loop marked “J” in the previous post) induced by the magnetic fields increases the electrical conductivity more effectively than at lower altitudes.
3. The more deceleration that can be done high up in the atmosphere, the lower the peak heating and the lower the total heat load.  The heat flux is roughly proportional to the cube of the velocity.
4. The heat flux reduction from this scheme is dominated by the increased shock layer thickness at high altitudes, and at lower altitudes is dominated by the much lower velocity by the time you get there by getting extra braking high up.
5. Conductivity of the plasma is one of the keys to making this work.  The conductivity in these cases was entirely due to the temperature in the plasma–higher velocities lead to higher temperatures, and Joule heating also leads to higher temperatures.  As velocities slow down, conductivity drops, as does the effectiveness of the braking system.  Below about Mach 12, the only way to keep the flow ionized enough to control magnetically is to add energy via some mechanism.
6. Because of the large induced currents, this idea only works if the heat shield is an electrical insulator.  If it is a conductor, you’ll just generate hall currents in the heat shield which will null out a lot of the benefit of the approach.

Thoughts on Maximizing the Effectiveness of MHD Reentry TPS
Based on these takeaways, and the discussion in the last post, I’ve come up with a few ideas for how to maximize the effectiveness of an MHD heat shield.

1. Use a lifting reentry.  Just as it is possible to offset the CG of a reentry body to generate some aerodynamic lift, it may also be possible to locate and orient the magnet in a way to create both lift and drag.  If you do a force balance on a body in a circular orbit, the downward gravitational force is exactly balanced out by a fictitious centrifugal force due to your forward velocity.  As you decelerate though, that centrifugal force component decreases, but by using lift, you can counteract some of that gravitational force.  This allows you to stay up at a higher altitude longer, which allows you to do more of your deceleration in the lower density air.  This is already used by all manned space capsules as well as the shuttle in order to keep reentry decelerations to a reasonably low level, and also to reduce the peak heating.  This is even more beneficial for magnetic braking concepts, because you can do more of your deceleration at a point where the magnetic effects dominate, electrical conductivities are high, and heat fluxes are low.
2. Use as strong of a magnet as you can reasonably work with.  While there are diminishing returns according to all of these papers, a stronger magnet does help provide more deceleration and shoves the boundary layer away further.
3. Use an alkali seed.  As velocities decrease, it gets harder and harder to maintain the electrical conductivity in the plasma at a high enough value to maintain useful levels of Lorentz interaction.  This is similar to the challenge with MHD electric generators.  In order to keep the conductivity high, injecting an alkali metal into the stream can help.  Alkali metals, particularly Potassium and Cesium have very low ionization energies compared to air.  In a weakly ionized plasma, most of the atoms are actually not atomized–almost all of the conductivity is provided by the small number of atoms that are.  So, a little bit of seeding can go a long way.  This helps you keep your magnetic deceleration forces high even as altitude and velocity drop.  The other nice thing about seeding, is that depending on what the fluid is, it might also cut down on the radiative heat transfer from the hot shock layer back to the heat shield.
4. Heat the plasma.  This may sound counterintuitive, but you might actually get better thermal protection if you start heating the plasma once you get to a certain point.  Below Mach 12, even with seeding, there just isn’t enough heat rise caused by the shock layer to keep the plasma  sufficiently ionized.  But, it is actually possible via several different means to dump a bit of energy back into the shock layer to push the gas back into an ionized state.  It’s unclear at this point if this is worth doing, but if the system is light and simple enough it might be worth considering.  As it is, you’ll have a lot of stored energy in the superconducting magnet, and you probably want to dump that somehow before landing–using it to keep the incoming air ionized a bit longer to get a little more deceleration before you hit the thick air might be worth it.

All told, you’re still going to need some sort of thermal protection for the last bit of deceleration, but the heat loads and max temperatures are so much lower if you can dump say half the reentry velocity while you’re still high up, that the problem becomes a lot easier to deal with.  If you could only get down to Mach 12 with this system, that would cut the peak and total heat loads by at least a factor of 8x.  The heat fluxes at this point would be low enough that you wouldn’t need ablative materials, and could probably use a ceramic tough enough that it was low maintenance.

Anyhow, the key questions I have at this point are: a) what sort of effective “L/D” ratio can you get by varying the location and orientation of the magnet, b) how much does seeding help, c) how long can you stay up in the high altitudes, d) what is the maximum amount of velocity decrease you can provide via this method, e) how strong of a magnet could you reasonably hold on an RLV, f) how does the strong magnetic field interact with the operation of the RLV itself–what does it do to solenoid valves, electric actuators, etc. and is there a way to shield against these issues?

In the next segments, I’m going to talk about another, possibly even more interesting application of this concept, as well as some thoughts on how we can reduce this technology to practice.

Oh, and one small caveat before I jump in–while I think there’s some real potential here, electromagnetics is a topic that I’m truly awful at. I’ve never had another class, including a PhD level turbulent fluid dynamics class that made me feel like such a brow-dragging neanderthal as my Physics 122 class on Electromagnetism. This may be yet another niche technology that while somewhat interesting, ends up not being all that useful. But it looks at least possible that this may become a game changing technology in many space transportation fields. Without further ado, let’s go over some of the basics.

Some Background on MHD Aerobraking and Thermal Protection
The basic concept behind MHD Thermal Protection is that during hypersonic flight, above about Mach 12, the shockwave formed in front of a blunt-bodied vehicle reaches a high enough temperature to form a weakly ionized plasma that is conductive enough to be manipulated by strong magnetic fields. A powerful magnet near the leading part of the vehicle interacts with charged particles in the plasma via the Lorentz force. This force bends the trajectory of charged particles, creates large hall currents, which if I’m understanding correctly repel the magnetic field. These charged particles also impact with the uncharged gas particles nearby (the plasma is only “weakly ionized”) transmitting these forces to them as well. Here’s an interesting diagram I’ll reference from one of the papers I’ll talk about more later (“Trajectory Analysis of Electromagnetic Aerobraking Flight Based on Rarefied Flow Analysis” by Otsu, Katsurayama, and Abe–well worth the $28):

Figure 1 (from Otsu et al): Schematic View of the Flow Around a Vehicle With Applied Magnetic Field and Induced Current

If the magnet is strong enough, this leads to two interesting effects–first, the distance from the vehicle to the bow shock increases (I think the plasma density between the bow shock and the vehicle also decreases, but I’m less sure about that). This can significantly reduce the heat transferred into the vehicle for a given velocity and altitude. The other big effect is that the Lorentz forces create forces that can produce drag or lift. At high altitudes these Lorentz forces can greatly augment the aerodynamic drag forces, effectively making it as though the vehicle had a much lower ballistic coefficient. It should be noted that this force is electrically controllable. In fact, depending on the sophistication of the magnetic apparatus and its location within and orientation with respect to the vehicle, it can possibly also produce lift as well as control torques without the need for aero control surfaces.

Both of these help from a reentry thermal standpoint, because by the time you hit the denser air, where the heating is the highest, you’re going a lot slower than you would’ve been otherwise, and a lot of that earlier braking is done at much lower heating loads than would have been the case without the electromagnetic effects.

Several of the papers I’ve read introduce an interaction parameter term, Q, that relates the relative strength of the Lorentz forces to drag forces. The relationship takes the form:

Equation 1 (from Otsu et al)

Sigma is the conductivity of the weakly ionized plasma, B is the magnetic field strength, L is a reference length (I think related to the magnet configuration), rho is atmospheric density, and V is velocity. As you can see, for a given magnet, the drag forces start dominating as the conductivity drops and as the atmospheric density increases. Atmospheric density increases dramatically as you descend from orbit, so for a reentry application, you get most of your benefit from the first little bit of descent.

We’ll go more into some of these ramifications starting in my next installment.

I’ve been a bit busy to weigh in on this for a while, but being a propulsion engineer myself, I can see *several* potentially interesting technologies that could be invested in.  To me, I don’t think the reason they’re suggesting investing in HLV technologies instead of building an HLV right now has anything to do with whether or not we can build one with existing technology.  My guesses at the reasoning are more along the lines of:

1-If they build an HLV right now and keep Orion going, they would either need to greatly increase the budget, stretch timelines out so far that it would drive the costs way up, or sacrifice the other parts of the plan that are what they’re actually interested in (tech demonstrators, ISS utilization, commercial crew maturation).
2-If they punt on actually building an HLV until after this first five-year chunk, then they’re no longer conflicting with the most immediate tech demonstrators or the commercial crew development efforts, which would allow them to build such vehicles without requiring a big increase in NASA’s budget.
3-By punting for a while, they both may have better technology options to start with, allowing for a cheaper, lower manpower HLV, but there will also be a much better baseline to start with for a new Orion vehicle.  By that point there will likely be at least two commercial capsule providers out there, and possibly some new tricks up their sleeve, which may make Orion a much smoother, better vehicle overall.
4-By punting on the actual HLV development until after the technology demonstrations, it may be clearer what capabilities you really need for the HLV.  Going ahead with one right now either has to assume that the tech demonstration will work (risky) or assume it won’t work (risky via overconservatism).  Waiting until you know more about if depots and high-Isp in-space propulsion systems are reasonable allows you to pick an HLV better matched to the new technology.

Going back to the technologies, if you understand that this isn’t a question of “we can’t do it with the technology we have today”, but more of a “we don’t want to do an HLV for other reasons, but what stuff could we do to make an HLV more affordable when we do have the budget to move on it”, then it becomes a lot clearer.  To me, the goal of any of the booster engine R&D should be to take technologies that have potential for drastically better cost/performance ratios, and mature them to the point that they could be rolled into an HLV effort when it is time to move.  TAN is one option, since if done right you could get enough thrust that a shuttle derived HLV wouldn’t need SRBs anymore to takeoff, and commercial EELV-class vehicles could either reduce their engine size, or increase their propellant load drastically for the same weight booster engine.  It might also allow you to get good performance with a lower pressure booster engine.  Other options include stuff like pistonless pumps, or some of the other pump concepts I’ve seen that while they may not have quite the performance of staged combustion, have most of the performance at a tiny fraction of the complexity/cost.  Flow separation control is another cool trick that can allow you to get more performance out of lower chamber pressure engines.  Imagine being able to build an engine that had better T/W ratio than an existing staged combustion engine, had no turbomachinery, and similar or better mission averaged Isp, but cost about a tenth as much?  That would be a game-changing set of technologies in my book.

There may be some other work relating to getting domestic high-thrust LOX/hydrocarbon engines ready.  There’s also the high thrust expander cycles you mentioned, which would be beneficial to EELVs as well as HLVs.  Depending on the engine size, it might even be interesting to further develop the Mid-Air Recovery idea that LM was investigating for their Atlas V engines.  Being able to recover the engines from an HLV launch and reuse them a few times without having to deal with salt-water corrosion issues is a great way to reduce some of the big marginal costs of a flight.  For upper stages, putting some money into ACES/Raptor related technologies might not be a bad investment either–find something that can give you some commonality between your HLVs and other stages flying (or make it so your HLV is just a family member of an upgraded EELV-class launcher family).  Or doing at least a little funding on reentry technologies could also be interesting (to allow for easier reuse of currently expendable booster stages, low the hurdle for true-RLVs, and also possibly make Orion-like vehicles easier to build and more capable down the road).  Some concepts like transpiration cooling and electromagnetic TPS/aerobraking are both really interesting.

Anyway, I’m an ideas guy and a propulsion engineer, so I may have a different view of this than most, but I for one think this is a reasonable approach.  To me, cutting back drastically on the amount of people who are needed to launch an HLV, while also using new technologies that allow you to get better system-level performance out of lower-cost, lower-complexity hardware is potentially game changing in my book.

To me one of the more interesting points is found at the top of page 11.  There were several misleading statements made by several people today about the relative safety of Ares-I compared to commercial crew vehicles.  As Brett put it (my emphasis added):

Second, some have claimed that NASA’s Exploration Systems Architecture Study (ESAS) shows that the current exploration vehicles are safer than commercial crew vehicles. In actuality, commercial crew vehicles were never even analyzed in the ESAS report – the ESAS report only looked at vehicles large enough to carry Orion, such as Ares I and variants of the triple-core Delta IV Heavy, and did not examine the smaller, simple, single-core vehicles, such as Atlas V Medium and Falcon 9 Medium that are sufficiently sized for commercial crew missions.  Moreover, even if ESAS had compared exploration vehicles to commercial crew-sized vehicles, the comparisons would be “apples vs. oranges,” because of the dramatically different tasks of these two types of vehicles.

When Jeff Hanley talks about how the Great Oz and supercomputers at NASA show that Ares-I is 3x safer than commercial launch vehicles, I wonder if he’s ever going to release their analyses for actually commercial crew vehicles, or if he’s being accidentally or intentionally dishonest. Because so far we haven’t been shown any data about the safety of actual commercial crew launchers. So far we have lots of data shown for the risks of using existing or modified commercial launch vehicles for launching a massive spacecraft designed to go to and return from the moon, including significant plane change maneuvers to allow anytime returns (ie Orion). It’s interesting to note that over half of the mass on Orion is the oversized launch escape system needed to get away from an SRB you can’t shutoff, and enough propellant for about 1500m/s of maneuvering to reach orbit and then to do in-space ops. That’s above and beyond the RCS propellant on the CM itself.

Most of the stuff that make Orion so massive are flat-out completely unnecessary for an earth-to-LEO crew capsule. You don’t need those kinds of delta-V capabilities. You don’t need as roomy of facilities, since by definition the flight times should be a lot shorter. Etc. There’s a reason why almost all of the proposed commercial crew systems are able to utilize single-stick launchers like Atlas V or Falcon 9–for an actual earth-to-LEO capsule you really don’t need anything bigger.

This realization that earth to LEO capsules can be much smaller than Orion leads to at least two important corollaries that I can think of:

1. Smaller capsules mean higher structural margins.  One of the existing vehicles most often suggested for commercial crew, Atlas V, was designed for the worst-case loading environment of any of its configurations (in this case I believe that would be the Atlas V 551 or 552).  The Atlas V 552 sees much higher max-Q’s than the 401/402 do, and has a much heavier payload on top, which exerts much larger structural loads on the Centaur stage than are seen in the 401/402 configuration.  While the Centaur structures may not meet the 1.4 magic number NASA likes in some of the bigger configurations, as I understand it, it actually exceeds that number in the 401/402 config most likely used for commercial applications.  The Falcon 9 was designed from the start to meet NASA structural margin specs.
2. No need for strapons.  Only one of the commercial crew ideas I’ve seen so far used a vehicle with strapons for crew launch (Dreamchaser).  This alone should make a huge difference in launcher reliability, since there are less things that can go wrong, less staging events, etc.  Most of the commercial launcher ideas they mentioned in ESAS assumed multi-core configurations.

There’s also the possiblity on the Atlas-V of using a dual-engine Centaur configuration to allow for some upper stage engine-out capability, or running the RL-10 at a derated performance level (not sure if that’s something it can do automatically, or if you’d have to make modifications–if you have to modify it it probably isn’t worth it).  With the much lower max-Q, and the ability to shut off the booster engine in case of an abort, I have a hard time believing that Ares-I is really that much more reliable than an actual commercial crew capsule launched on a commercial launch vehicle that has dozens of flights under its belt.

Food for thought.