[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!

All that aside, this post is related to Joe Carroll’s last topic–reduced gravity research facilities. His talk reminded me that I needed to dig-up and finish this blog post I had started back in May about the importance of such reduced gravity research facilities, and a clever approach I had seen to providing them.

Reduced Gravity Effects on the Human Body
I first made this point almost five years ago, but it bears repeating: while we have a lot of data on human health at 1g and at 0g, we have almost no data in the middle. I say almost, because we did have a dozen people live on the moon for at least 24 hours each…but that’s pretty much the only data we have on reduced gravity health effects, which is far too little to draw any really useful conclusions.

Most readers of this blog know that the data from microgravity impacts on the human body don’t look too promising–even with lots of exercise, there are apparently biophysical mechanisms that can have large negative health impacts (osteoporosis, neurological and pulmonary issues, etc) that begin to show themselves very quickly. However, as I pointed out in that earlier post, we have no idea which of the curves below really represents human health impacts of reduced gravity:

Does just a little bit of gravity go a long way (my personal guess I explain in the other post)? Or do you need almost full earth gravity? Or is there actually some gravity level less than 1g that’s actually better than earth gravity? While natural selection for humans has obviously been focused on a 1g environment, that doesn’t mean that humans are so hyperoptimized to 1g that nothing else will do. It’s unfortunately possible, but right now we don’t know. Without getting some “center points”, any guess at the shape of the response curve is just that–a guess.

Why This Matters
The reason why this knowledge void matters is that it greatly impacts the future expanse of humanity into space, as well as near-term human exploration. For instance, we don’t know if someone who goes to live on the Moon or Mars can ever really come back to earth, or if they have kids, if their kids can return.

If however it turns out that lunar gravity is already enough to counteract the worst of the effects of microgravity, it might be that the best way to do initial lunar human exploration is something like a One-Way To Stay (for a while) approach. If you knew that you could send someone for long durations while still being able to bring them back later if needed, it would open up some big possibilities. The return portion of a human lunar mission is one of the big performance drivers that make human missions so much more expensive than robotic ones. Even if you couldn’t close the life-support loop, just not having to return the initial explorers right away could allow you really enhance robotic exploration of the Moon by having people there on the spot to help troubleshoot, fix, upgrade, iterate, etc on your robotic systems. I know a lot of people think we can just send robots and have them make a turn-key base. It’s possible, but I expect you’re going to break a lot of robots along the way, and you could avoid that by having people in the loop. But its ethically hard to do a mission like that before you have some data on what long-duration exposure to 1/6g is going to do to your explorers.

Returning to the Joe’s talk, he suggested looking at .06g as well as lunar and martian gravity, as a possible minimal gravity level that people could intuitively adapt to without lots of training. If travelers can get by without large negative health hazards by .06g worth of gravity, that would really simplify the concept of providing artificial gravity for long-duration deep-space trips (like to Mars or NEOs). If there’s a “knee in the curve” above which you can avoid the worst of microgravity effects, that can make it a lot easier to provide artificial gravity for trips like that. If you have to provide a full 1g, and can’t go with high RPMs (which Joe suggested that the terrestrial centrifuge data might be suspect due to the presence of a 1g downward gravity vector), that implies very large structures, which become a much bigger engineering challenge.

xGRF
The question becomes, what’s the best way to get this data? Most of these effects take timescales on the order of hours, days, or weeks to express themselves. And there’s no way on earth to adequately simulate hypogravity. The only real way of testing this, short of going there and finding out the hard way, is to build some sort of orbital research facility. The ISS was originally going to have a Centrifuge Accommodations Module, but that project got defunded, and the hardware is no longer flightworthy from what I hear. I had suggested the idea of doing a “CAM in a Can” before, but even that would be limited to studying small animals–there’s no way you could fit a human in there. To get the data quickly, you really want some sort of artificial gravity facility that is human-sized. In his presentation, Joe Carroll talked about building a large rotating space station with facilities on different lever arms from the CG of the facility. While this is interesting, and would allow you to have your gravity decoupled from your spin rate, I think that Kirk Sorensen’s xGRF “Variable Gravity Research Facility” concept makes more near-term sense (Joe and I disagree on this point BTW).

I’m not sure if Kirk reads this blog very much anymore (he’s pretty busy at his new job as Chief Nuclear Technologist at Teledyne Brown), but I have to toot his horn a bit. While not all of his ideas are ones I’m sold on, he’s had more than his fair share of clever ideas. The idea behind xGRF is very simple. You have a small facility–something on the scale of a Sundancer or Nautilus module from Bigelow, and you attach it via a long tether to a large counterweight (such as the upper stage that delivered the module to orbit in the first place). In LEO the gravity gradient can be used to force such a system to adapt an orientation with the long axis pointing through the center of the earth. In such a situation, the CG will be somewhere between the two end pieces, and the module will be going slightly slower than the orbital velocity of other components at its altitude, and the counterweight will be going slightly faster. This provides a tiny bit of settling force on each end (acting like a tiny bit of gravity with a vector pointed outward from the center of the system).

Ok, you may be thinking, that’s nice. But where do you get the “Variable” Gravity from? That’s where Kirk’s idea gets really clever.

Basically, something in a gravity gradient orientation is still actually spinning–it just completes one complete rotation per orbit around the earth…What happens if you take a spinning object like this, and decrease it’s moment of inertia by, oh say winching in the tether? By conservation of angular momentum, the object has to start spinning faster!

You can winch the habitat and the counterweight together until you reach the desired level of artificial gravity. Depending on the design details, you can pick any gravity level you want between say microgravity and 1g. How do you dock, say to transfer crews or deliver supplies? Well, it turns out you can despin the system by just reeling out the tether:

Pretty clever. By doing this, not only can you pick any gravity level you want, but you can also do your rendezvous and docking in a simple, non-spinning environment, you can eliminate the need for having rotating and nonrotating parts of the station, or of long elevators or connecting tunnels. I really like this concept, because the system ends up being pretty simple, with everything being able to be launched on a single EELV flight. You don’t have to assemble a huge space facility and then spin it up. This can be a small project that might actually get built. I think the big station Joe might have more capabilities, but I’m worried that detractors would paint it as a second ISS, and it would never get funded. Something on this scale though is within the realm of feasibility.

Flagship Technology Demonstrators, Expansion Options, Future Uses, and other Parting Shots
One particularly interesting way to get something like this funded (and what I was originally writing this blog post back in May as a response to) is as a replacement for the “Inflatable Technology” Flagship Technology Demonstrator. Back in Galveston late last spring, NASA rolled out several proposed FTD missions to flesh out plans suggested in Obama’s FY11 budget proposal. One of the missions was to build an inflatable module and attach it to ISS. To be honest, this seemed a little duplicative–it looked for all intents and purposes as though NASA was going to spend $500M-1B duplicating what Bigelow was doing on his own dime. I think a much better way of both flight demonstrating inflatables while killing multiple birds with one stone would be to build something like xGRF as a Flagship Technology Demonstrator. Leverage either a Bigelow Sundancer module or compete it out and have ILC Dover also bid on it. For the same amount of money, you get a much more useful lab, that doesn’t endanger the ISS, and which allows you to do reduced gravity research that compliments ISS’s microgravity focus.

As Joe pointed out, even after the initial experiments (say at lunar gravity first, then Martian, then at the .06g level), a facility like this would have lots of follow-on utility. You can answer initial questions relatively quickly–ie even a few months at each level would tell you a lot compared to what we know right now, but getting longer-duration data could be very useful for future space settlement efforts. I’ll have to dig up my notes on all the reasons, but there’s a lot of long-term potential for a station like this.

Which means you might also want to upgrade it down the road. If you overbuild the tether, and the docking facilities, you could probably attach additional modules to a station like this pretty readily. To add to the counterweight, you could say have facilities on the original upper stage that could allow it to be outfitted as a depot…but that’s getting a little too crazy for now.

But I think the time for something like this is now. FTDs are getting money, even if it’s greatly reduced from what Obama wanted. The budget for exploration technology development, including flagship missions is currently authorized at over $1.1B over the next three years. At that rate, you could fund most of the work on both the depot approach that was proposed by the joint industry/NASA group I participated in last year, as well as xGRF, and still have money left over for starting another FTD like say an aerobraking or aerocapture one. Even if funding gets further reduced in appropriations, there’s enough money to pursue something like xGRF and depots in parallel.

I think this is an idea whose time has come.

Anyhow, the papers referenced in Valterri’s article were fascinating. I really would suggest reading the whole thing if you have the time. While it’s important to take such early reports with an appropriate sized grain of salt, I think this could be an interesting avenue of research for NASA to fund. I’ll definitely be keeping an eye on this research now that it’s been brought to my attention.

[Update: In case you guys didn't guess, this was an April Fool's joke. Don't feel bad if you got snookered, Valterri had me going for about 30min last night. I only realized it was a joke when I saw the date stamp. April Fool's starts early over there in Europe...darned timezones...]

Traditional Aerobraking and Aerocapture
One of the challenges of orbital mechanics is that it takes just as much energy to descend into a gravity well as it does to ascend out of it. One technique that has been used for lowering the propellant cost of descent into the gravity well of a planet with an atmosphere is aerobraking. Aerobraking is the process of taking a spacecraft in an ellpitical orbit around a planet with an atmosphere, and using atmospheric drag at the lowest altitude portion of its trajectory to slowly decrease the altitude of the high end of the elliptical orbit. This process has been used now on about a half-dozen planetary missions, in some cases reducing the propulsion requirements by 1km/s or more, over the course of a couple hundred passes. Aerobraking has been traditionally been done by satellites that aren’t explicitly shaped like a reentry vehicle–in fact most of the drag for typical aerobraking vehicles is produced by using the spacecraft’s solar panels as massive drag brakes!

Artist's Impression of MRO Aerobraking (credit JPL and Wikipedia)

A more aggressive maneuver called aerocapture takes a spacecraft in a hyperbolic (interplanetary) orbit and in a single pass decelerates that vehicle into an elliptical orbit around a planetary body.  Typically the term refers to maneuvers where the ending orbit has an apoasis near the altitude of a circular orbit, though it could also be used to describe a maneuver that uses a single pass through the atmosphere to replace the “capture braking burn” that would normally be used. Aerocapture is a lot more challenging, since the deceleration has to take place a lot lower in the atmosphere in order to provide the required deceleration in such a short distance. This implies much higher forces and heat-fluxes, which require some sort of aeroshield/TPS system.

Here are a few of the main challenges of aerobraking and aerocapture:

1. Dynamic Pressure Loads: Dynamic pressure is the pressure felt on the vehicle by the impingement of the atmospheric molecules.  The equation for dynamic pressure is q = 1/2 * rho * V^2, where lower case q is the dynamic pressure, rho is the instantaneous atmospheric density, and V is the instantaneous relative velocity.  For MRO, the dynamic pressure limits were set at 0.35 Pascals, which correlates to moving at about .76m/s at sea level (ie a slow walking pace).  To give you an idea of how this compares with orbital reentrythe peak dynamic pressure of say a Soyuz in its emergency ballistic reentry mode, is over 40,000 Pa of dynamic pressure, and even a low-G lifting reentry is still in the 10kPa+ range.  Direct entry into the Venusian atmosphere from a hyperbolic interplanetary orbit gets you into the 1MPa range!  Another fun comparison is that the max-Q Xombie or Xoie have seen in flight was around 250Pa.Most of the very low allowable dynamic pressure load for past aerobraking efforts has been driven by the fact that most aerobraking craft to-date have used large flimsy solar panels as their main drag structure.
2. Peak Heat Flux:  The shockwave caused by slamming into gas particles at hypersonic velocities compresses and heats the gas particles to substantial temperatures.    Heat from this shock wave is convected and radiated into the aerobraking spacecraft.  The equation for heat flux is Q = 1/2 * rho * Ch * V^3.  Capital Q is the heat flux (in W/m^2), rho and V are the same as before, and Ch is the heat transfer coefficient.  The heat transfer coefficient, I think, represents what portion of that heating goes into the vehicle itself instead of being carried off by the now quite ruffled atmospheric gas molecules who didn’t see you coming.  Yes it is confusing that dynamic pressure is lower-case q, and heat flux is capital Q.Once again, to give you some scale, the worst case pass for Odyssey had an estimated heat flux of about 500 W/m^2,  which is about 40% of the heat you get in LEO from the solar radiation. For that Soyuz reentry case mentioned earlier, the total heat generated at max-q is in the 240 MW/m^2 range–several times higher than the heat flux at the throat of the SSME or RD-180.  The Venusian direct entry example according to one source would actually be in the 4000MW/m^2 range! Fortunately, I think that for atmospheric reentry the Ch term is relatively low–most of that heat gets carried away by the atmosphere.As with dynamic pressure loads, the reason why peak heating rates are kept so low for most aerobraking missions is that you’re using the large solar panels as most of the drag surface, and they can only take so much heating before their temperatures rise to levels that could permanently degrade their performance.
3. Atmospheric Density Variations: If atmospheric density was nice, constant, and well-known, aerobraking could proceed a lot faster and in a lot fewer passes.  The problem is that at the altitudes where aerobraking takes place (100+km), the density can vary significantly over length scales as small as 20km.  This can be driven by many processes including variations in the solar wind and solar radiation due to sun cycles, weather effects like dust storms for Mars aerobraking, and other effects.  Going off of some data from the Odyssey mission, variations as big as 2-3x were seen in density from pass to pass.   A second-order effect of density variations is that both the drag coefficient and the heat transfer coefficient will vary with atmospheric conditions by noticeable amounts.  Unfortunately,  in many cases you don’t know the density along a given trajectory in advance, so you have to plan for not the average density, but the worst case pass density.   Which means that most of the time you’re getting less deceleration and heating than you could actually withstand, but some of the times you might actually find yourself pushing your limits more than you would like.   This drives you to taking more passes than you’d really like to take in an ideal situation.  These variations get more and more pronounced at higher aerobraking altitudes, where atmospheric density is measured in kilograms per cubic kilometer.Once again, this is an area where using large, sensitive solar panels as your drag devices really hurts. Because you can’t stand high dynamic pressures or heat fluxes, you have to do your passes higher up in the atmosphere. But due to variability in density at those higher altitudes, you end up getting driven even further up to deal with worst case variations. That said, even aerocapture trajectories are high enough altitude that atmospheric variations can be important challenges to deal with.
4. Aerobraking Duration: For most previous Mars and Venus aerobraking missions, velocity changes in the 1-1.2km/s range have taken between 70-150 days, over several hundred passes.  While this is fine for unmanned missions, it’s harder to do for manned missions, where radiation concerns make you want to minimize your time spent in-transit.  The large number of cycles is also a difficulty for missions aerobraking at earth, where each pass will take you through the Van Allen belts.  Lastly, for reusable in-space transports, the total turn-time is an important economic parameter–the more missions you can fly in the same period of time, the fewer vehicles you need to support a given mass throughput.

A couple more quick observations before we jump into using MHD forces to enhance aerobraking:

• For typical aerobraking, the parameter you can control easiest is the periapsis altitude, and thus indirectly the average density.  In other words, if you want to double the drag on a pass, you lower your periapsis to an altitude that has about double the average density.  This also means that to a first order approximation (ie ignoring the relation between density and the heat transfer coefficient) heat flux for traditional aerobraking is going to scale fairly linearly with drag.
• Ballistic coefficient ends up being really important for aerobraking as well–this is the whole reason why the solar panels are used unstowed for aerobraking.  Higher ballistic coefficients mean that you have to dip lower into the atmosphere (and thus get a higher heat flux) to get the same amount of deceleration per pass.
• In spite of the disadvantages of using solar panels as your drag brakes, there are some real advantages to being able to use a aerobraking scheme that doesn’t require your vehicle to be explicitly crammed into a typically reentry-vehicle shape behind a massive heat shield.  It would be nice for instance to be able to get tanker vehicles or orbital tugs back from lunar trajectories or martian trajectories without them having to carry a big aerobraking shield like you see in all the old literature.

Anyhow, that was a quick introduction to aerobraking by a complete non-expert.

Some Backstory on Why I’m Interested in Aerobraking
I started looking into this a few months ago as an alternative to propulsive retrobraking for Centaur-derived cislunar tanker vehicles.  While a Centaur stage actually can do a lunar round trip fully propulsively, with at least some payload delivered to the Moon, the “gearing ratio” (initial mass in LEO compared to payload delivered to LUNO or the Lunar Surface) was pretty pathetic.  Just to use some ballpark numbers, without digging up my more precise calculations, I’m getting around 8000lbs payload to LUNO if you drop it off in orbit and the Centaur only returns to earth, dropping to only 2500lb if the Centaur has to haul the payload all the way there and all the way back propulsively.  However, if you could do 3km/s worth of aerobraking (assuming about 1200m/s worth of burns between the Trans-Earth Injection burn and any periapsis raising maneuvers, including the final circularization), all of the sudden you’re talking about almost 20,000lb of payload on the dropoff mission, and about 13000lb on the round-trip maneuver.  Depending on how massive and expensive the aerobraking system weighs, it makes a massive difference in the performance of a reusable cis-lunar architecture.  For a long time though, I had sort of dismissed aerobraking, because any aeroshield big enough to allow single-pass aerobraking (or few enough passes to be interesting) also ended up looking like it would either be very heavy, or very bulky, or require lots of orbital assembly or some sort of new deployable technology.  Not that any of those other than being too heavy was a total show-stopper, but it definitely made it less attractive for a near-term commercial operation.

Another line of thought I had been wondering about recently was manned cislunar transportation, especially in light of the Augustine Committee report.  One of the big suggestions they made that rubbed a lot of HLV-advocates wrong was the idea of launching the crew on commercial LEO taxi vehicles, and flying Orion up to LEO unmanned.  A lot of people said this was just silly–if you’re launching Orion may as well launch it manned, even though this would require adding launch escape and emergency detection capabilities to the HLV.  I started thinking down the lines of what Orion could look like if it was designed from the start not to carry astronauts until they got to space.  The LAS would go away, as would all the structural requirements for taking those sorts of loads, being able to rapidly drop the service module, etc.  The whole thing could fit inside a fairing, thus simplifying aerodynamics and loads on the front end of Orion.  Heck, it could even be attached to the rest of the stack in whatever orientation made the most sense for mission ops–it wouldn’t be constrained by needing to be on the top in an orientation where the capsule could “get out of Dodge” in a hurry if something “went south” with the HLV.  The more I thought about it, the more I realized that Orion could end up looking like a drastically different vehicle if it was optimized for in-space use and reentry instead of needing to also handle manned ascent to orbit as well.  Then I made an interesting leap of logic.  What if Orion was only meant to be used in space?  I originally sort of dismissed this, since most single-pass aerobraking schemes I knew of would require the thing to be designed like a reentry capsule anyway.

Jumping back to the Centaur-based tug idea, I toyed around with the idea of doing a blog series, seeing if I could make an aerobraking simulator to figure out if a Centaur could without any sort of fancy aerobraking shield actually do a multi-pass aerobraking mission that would get it back to LEO within a reasonable amount of time (say three weeks or less).  However, I stumbled on the papers about magnetic aerobraking right about this point in my thought process, which may possibly provide a solution to both of these problems.

While I don’t have anywhere near the analytical chops to know for sure how far you can push this technology, if it could enable single-pass or at least small number of pass aerobraking without requiring a huge traditional aerobraking shield, interesting things might become possible. Magnetic aerobraking could potentially revolutionize cislunar transportation, enabling low-cost reusable manned and unmanned deliveries based on modified versions of existing LOX/LH2 upper stages, and could allow fully-reusable in-space only manned vehicles that weren’t just an overglorified 1960s-style reentry capsules.  But more on that later.

For now let’s get back to how we can use magneto-hydrodynamic interactions to enhance traditional aerobraking, and see if we can figure out if this idea has merit at all.

Magnetic Aerobraking
Going back to our previous two discussions, one of the key takeaways was that the enhanced braking and thermal protection provided by strong magnetic fields was strongest at high altitudes where atmospheric density was lowest. At high altitudes, the ambient atmospheric density is low, but Joule heating caused by the interactions between ions in the shock layer and the superconducting magnet keeps the electrical conductivity of the plasma in the shock layer high. Also, for aerobraking or aerocapture short of reentry, by definition you are both always at a speed and altitude high enough that you don’t have to worry about the shock layer losing sufficient conductivity for MHD effects to dominate aerodynamic drag effects. The magnetic interaction parameter (Qmhd) introduced in my first post in this series can easily be in the 250-1000+ range at high altitudes compared to down in the 5-50 range you might see during atmospheric reentry. For example, the paper I cited in my first article (Otsu et al) showed that for a vehicle coming back from a GTO-like orbit, you could cut the return time by 70% with a 0.1T magnet, which is about 5x weaker than the magnet assumed for most of the reentry magnetic TPS studies.   While magnetic effects may be helpful for reentry, they truly come into their own for aerobraking and aerocapture.

A few other thoughts:

• While the total drag for a magnetic aerobraking concept can actually be several times the drag of a similar non-magnetic vehicle, the gas-dynamic portion of the total drag actually decreases substantially in the case of magnetic aerobraking.  This is due to a much lower velocity behind the shock layer in the magnetic case.  Figure 9 from the Fujino et al paper I used in the last post (“Numerical Analysis of Reentry Trajectory Coupled with Magnetohydrodynamics Flow Control”, JS&R Vol 45 No 5, pg 911-920) illustrates this beautifully:
• For a vehicle using magnetic braking, most of the total drag force is actually reacted electromagnetically through the magnet itself, not through the surface of the vehicle.  The dynamic pressure that the vehicle surface itself sees is greatly reduced compared to what you would expect at that altitude and entry velocity.
• While in the above case, the dynamic pressure reduction was about 4x at ~75km, this effect is likely to be even more pronounced at the altitudes used for aerobraking (90-120km) where the electromagnetic interaction parameter is substantially higher (40-160x higher) than it is in the case shown above for atmospheric reentry.
• The heat flux seen by the aerobraking vehicle will also be greatly reduced compared to a non-magnetic aerobraking system at a similar altitude and velocity.  This is due to the much thicker shock layer standoff distance and the lower velocity of the particles behind the shock layer.  The Fujino et al paper estimated that the heat flux would roughly be cut in half at 75km with a 0.5T magnet (due to a boundary layer between the bow shock that is twice as thick at that magnetic interaction parameter).
• For higher parameters in the 100-1000 range that you would likely see for aerobraking, this effect should be even more pronounced.  The trend in shocklayer thickness vs. Qmhd shown in Fig 3 of Fujino et al  was linear over the Qmhd range of 0-6.  If it continued out linearly up into the Qmhd 100-1000 range, the shock layer standoff distance would be in the range of 100-125x thicker than without MHD effects, implying a drastically reduced heat flux at aerobraking altitudes.  Unfortunately without having them run the actual analysis, it would be hard to know precisely how well this would work.
• All these factors mean that the same vehicle could use a lower periapsis with a magnetic braking system than without.  The dynamic pressure and heat flux that the vehicle sees at a given periapsis altitude is going to be at least 2-4x and possibly more than an order of magnitude less than it would be without the magnetic field.  Even in the most conservative case (ie assuming that the effect at 100km and aerobraking speeds is no better than at 75km in spite of having a Q 40-160x higher) this would allow you to go to an altitude with at least double the density while keeping the heat flux and dynamic pressure loads within tolerances.  With an effective total drag 4x higher at a given altitude combined with being able to go to a lower periapsis, you get bare minimum a 8x reduction in total aerobraking time compared to the non-magnetic case.
• For the aggressive, “I don’t know if I’m extrapolating way too far” case, you could get even larger reductions in aerobraking time.  Going back to my linear extrapolation on shock layer standoff vs. Qmhd (and thus heat flux vs Qmhd), at Qmhd=250 this would put the shock layer standoff at about 25-30x thicker than the non-MHD case.  The example in Otsu et al gave a Qmhd of 250 using a 0.1T magnet and a 100km periapsis.  Since Qmhd is proportional to B^2 and inversely proportional to rho.  If you increased the magnetic field from 0.1 to 0.5T (similar to what was being suggested for the reentry studies done by Fujino et al and some of the others), you could maintain a Qmhd of 250 even if you increased the local density by a factor of 25.  At Qmhd of 250, the effective drag coefficient is about 3x higher than the non magnetic version.  That would give up to a 75x reduction in aerobraking time compared to the non-magnetic case.
• One other advantage of magnetic aerobraking is that you can drastically vary your effective drag coefficient electrically.  Also, the heating and dynamic pressure are far more driven by the magnetic field strength than by the atmospheric density for the MHD aerobraking case.  These mean that you can afford to take deeper passes without having to worry as much about variability.  If the density is higher than expected, and you have some head-room on your magnet, you can increase the MHD field strength a bit to keep the shock layer back and the dynamic pressure down.  This also could cut trip times in half just by allowing you to base your planning off of the average atmospheric density instead of having to take the mean + 3 standard deviations as your predicted atmospheric density.

I’m rapidly coming up to the point where I’m pretty sure I no longer know what I’m talking about.  At least from here, it looks like there’s a good chance that MHD aerobraking could allow for aerocapture (at least into a high eccentricity elliptical orbit), and very rapid aerobraking down to a circular orbit compared to the non-magnetic case. I think you can extrapolate the conclusions of these papers in these ways, but without having the people with the analysis tools actually verify these claims, I’d still take them with the appropriate sized grain of salt.   Also, my intuition on how a MHD aerobraking vehicle would compensate for density variations is not very good.  That alone could be a paper or a thesis.

So, whether this ends up being a mild curiosity that ends up only being useful in niche applications, or a game-change remains to be seen, but the potential for this being a game-change is real.

In my last post in this series, I’ll go more into some of the implications of what this could do if it works, and some thoughts on how to actually flight-demonstrate MHD aerobraking.

[Edit: It turns out I had misspelled Fujino's name in the original post. Fixed that and added the title of the paper in case people want to get a copy--it's free if you have a JS&R subscription, $15 if you're an AIAA member without a JS&R subscription, and something like $30 if you're not an AIAA member--highly recommended if you're interested in this topic]

The expression uses the Gaussian error function, erf(x), which is not typically available in a spreadsheet or scientific calculator. erf(x) also cannot be calculated in closed-form–typically the expressions used to calculate it are iterative. Since the tapered tether mass ratio is such a useful design tool to have, I derived a recursive algorithm that computes the ratio in a simple loop, given only the velocity ratio of the tether (tip velocity/characteristic velocity).

Here is the recursion:

and here is psuedocode for the recursion:

subroutine getRatio(double VR, int k) {
double VR2 = VR*VR;
double sum = 0.0;
for (int i = k; i >= 1; i--) {
sum = (VR2/(double)i)*(1.0/(double)(2*i+1) - sum);
}
return 2.0*exp(VR2)*VR2*(1.0 - sum);
}

With about 8 recursions, the results are extremely accurate. The recursion is unstable when the velocity ratio is greater than about 3, but no one should be building tethers with velocity ratios greater than 3! and you can just use Moravec’s expression with erf(x) = 1.0, which is a pretty safe assumption for x > 2.0 or so.

The first idea of concept of a tether dates back to the imagination of Konstantin Tsiolkovsky, the Russian schoolteacher who first developed our modern concepts of rocketry and first derived the rocket equation. In the late 1800s, Tsiolkovsky visited Paris and saw the Eiffel Tower. He was so impressed by the sight that he imagined a tower reaching up far into space. He calculated the height at which such a tower would have to be before the centrifugal force from the Earth’s rotation balanced the pull of gravity (inadvertently calculating the altitude of geosynchronous orbit).

Tsiolkovsky, of course, could not conceive of any material that could withstand the compressive forces of such a structure, but sixty years later, a Russian engineer named Yuri Artsutanov picked up the thread of Tsiolkovsky’s work and first worked out the engineering principles of what is now called a “space elevator”, a long tether hanging all the way from geosynchronous orbit to the surface of the Earth. The space elevator required materials with specific tensile strength far in excess of any known material, and still does. Further conceptual engineering work on the space elevator concept was done in the early 1970’s by American engineer Jerome Pearson.

The space elevator was a hanging tether, and payloads were required to traverse its length in order to achieve orbit. The beginnings of rotating momentum-exchange tethers date to the late 1970s, when Hans Moravec, a robotics researcher at Stanford University (now at Carnegie-Mellon) was intrigued by a suggestion of his friend John McCarthy of a satellite that “rolled like a wheel” around the Earth. Moravec began a scientific investigation of the concept, which he first called a “non-synchronous orbital skyhook” and later a “Rotovator”. Like the space elevator, it reached all the way to the surface of the Earth, but unlike the elevator, it rotated about its axis a number of times per orbit. A payload would be picked up by the tip at the surface of the Earth and then thrown half a rotation later into an interplanetary trajectory. The Rotovator was a good deal shorter than the space elevator (~4200 km vs. 40, 000 – 100,000 km) but was not much better in terms of materials required. Moravec published a paper on the subject in the Journal of Astronautical Sciences where he speculated on advanced forms of matter that might have the strength needed to build the Rotovator.

About a year after the JAS paper was published, Dupont’s development of Kevlar excited Moravec to the possibilities of Rotovators built with conventional materials. He wrote a short paper called on the subject which was never published. The paper showed that Kevlar skyhooks were not feasible around the Earth but could be reasonably built around the Moon and Mars. In an appendix to this unpublished paper, Moravec speculated on the possibility of skyhooks built in interplanetary space that would assist spacecraft traveling between the Earth and Mars. To the great benefit of future tether researchers, his equations for the cross-section of a tether, in the absence of a gravitational field, could be integrated in closed-form. Thus, the Moravec “tether equation” was first derived.



Moravec was able to derive analytical expressions for the area of the tether as a function of its distance from the rotational center. He then numerically integrated the area expression along the length of the tether to calculate volume and mass. As an aside, in an appendix, he considered the case of a tether spinning in free space. When the tension on the tether was only due to centrifugal forces, the area expression could be analytically integrated to a closed-form solution. Thus the Moravec mass ratio was derived.

The equation could be simplified by realizing that fundamentally, the mass ratio is a function only of the velocity ratio, which itself is the ratio of the tip velocity of the tether and the characteristic velocity of the tether material.


Further insight into the value of the equation was gained by comparing it to the rocket equation and noting the similarities and differences.

Moravec wrote a few articles on the subject for space-themed publications, but basically returned to his robotics work. Nevertheless, Moravec’s equation still serves as a foundation to all momentum-exchange tether work to this day.

The rendezvous required for a rotating tether and its payload is far more dramatic. The whole point of the operation is to have the tether and the payload in different orbits, so that the rendezvous can lead to an exchange in angular momentum and orbital energy between the two, resulting in a payload boosted to a higher energy orbit (or dropped to a lower energy one).

Thus, you can’t match orbits like you do in conventional rendezvous. The best that you can do is to instantaneously match position and velocity (but not acceleration). So you need an approach to rendezvous that is pretty tolerant of error.

So we threw out the book when it came to trying to think of how to do rendezvous, and came up with something totally different and designed to meet the specific needs of the mission. And I was pretty proud of the result, and still am. Because, you see, this is a bit of an anniversary for tether rendezvous technology. It was five years ago (February 2005) that we successfully demonstrated that the rendezvous technology we had postulated could work, at least at the lab scale.

We took advantage of the fact that the tether was under rotation and experiencing centrifugal acceleration, and that the payload was in free-fall. We simulated this (quite accurately) by hanging the tether’s “catch mechanism” from the ceiling of a racquetball court at Tennessee Tech, and then we “shot” our simulated payload up to the catch mechanism, with its boom positioned to be captured by the catch mechanism when it penetrated the aperture of the catch mechanism. Then the catch mechanism would release and close around the boom, quite quickly, allowing the simulated payload to be caught.

It all worked out a lot better than I thought it would–take a look at our results:

First Catch Mechanism Test

Second Catch Mechanism Test

More Testing with Animation

And here was the press release that came out months later announcing the accomplishment. Our video footage of successful testing got on NASA TV…once.

NASA Engineers, Tennessee College Students Successfully Demonstrate Catch Mechanism for Future Space Tether

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.