Carrier Craft By-the-Slice
Anyhow, the key thing that led to this concept was the realization that for air-launched vehicles, you really want to be able to buy the carrier aircraft “by-the-slice”, instead of having to own it outright.  Ie, you want to be able to call up Virgin Galactic and say “what’s your schedule for this week…ok, can I buy the noon-5pm slice on your WK2 next Thursday?  Usual ammenities.  We’ll meet you on the flight line at 12 o’clock sharp.  Thanks!”

Especially for early-generation airlaunched RLVs, not having to pay the full burdened cost of an aircraft in addition to the rocket stages is really important.   It’s even more important to avoid having to pay the development cost of a custom carrier craft.   If you only have to rent the carrier craft when you need it, you are much more able to handle ups and downs in demand, which will likely be fairly volatile for the first few years.  It also means that you can use the carrier plane during development without having to have that big of a capital investment early on.   IIRC, this was one of the keys to making Zero-G’s parabolic services possible–the same aircraft can be converted to cargo-hauling overnight.  I don’t know if they still use this capability, but not having to pay for the full aircraft just out of space revenues makes it easier to charge an attractive price.

The problem is that there aren’t any airlaunch carriers that you can “buy by the slice” that are big enough for an manned orbital RLV.  That’s where exo-atmospheric refueling could come in.

Exo-Atmospheric Suborbital Refueling
Many years ago, people in the military realized that being able to refuel aircraft in-flight could greatly expand their operational capabilities.  Of course, at the time that the idea had first been raised, mid-air refueling sounded about as crazy as exo-atmospheric refueling sounds today.  The first mid-air refueling involved linking two biplanes, and having a guy walk from one aircraft to the other carrying a 5lb gas can!  Of course, over time, much safer procedures and techniques were invented, and nowadays its easy to forget how insane this idea must have looked before it had been tried.

That said, while it might be possible in some cases to eliminate the need for mid-air refueling by using a bigger vehicle with larger propellant tanks, there are some missions that would be flat-out impossible without the capability.  The typical response I get when I suggest something crazy is, “why don’t you just build a bigger rocket”.  While there are some cases where you’d be better off just “building a bigger rocket”, there are some times where additional operational complexity more than pay for themselves.  And I think the added complexity in this case is worth it if it buys you the ability to use an off-the-shelf carrier plane, bought by-the-slice, and keeping your rocket stage sizes small, and the rocket engine sizes in the low 10s of klb range, while still being able to feasibly deliver people to orbit.

So, what do I mean by exoatmospheric suborbital refueling?

Basically, it means that at some time after the vehicle leaves the main sensible atmosphere, it hooks up with another vehicle on the same trajectory, propellant is transferred from one vehicle to the other, and then the first vehicle continues on to orbit (while the mostly-empty tanker vehicle reenters and lands). [Note: Gary Hudson reminded me that Mitchell Burnside-Clapp of Pioneer Rocketplane investigated just such an enhancement to their system, during their work on the RASCAL project.]

The configuration that I’ve used in my analysis is a pair of TSTO vehicles launched off of a WK2.  The first stage in both stacks is identical, and are roughly 25klb wet, 5klb dry.  The orbiter and tanker stages are about 10klb wet each, with the orbiter upper stage launching a little more than half full, and the tanker stage having smaller tanks fully-loaded on takeoff.   The tanker and orbital stage would be built to more aggressive mass fraction targets than the first stages. [Note: you can find a copy of the spreadsheet I used, in case it's helpful, here --Jon]

The operational procedure would be something like this:

1. Both WK2′s head uprange to the appropriate drop zone, spaced as close as is reasonably possible–maybe 1km apart?
2. The two TSTO stacks drop and light at the same time.
3. During the first stage burn, the two stages slowly close the relative distance between each other, so that at staging they’re maybe 50-100m apart.
4. As soon as the dynamic pressure is low enough (possibly even before staging), booms are extended between the two orbital stages.
5. The first stages separate from the upper stages, and glide forward from the staging point to the launch field for landing and reuse.
6. The upper stage and tanker stage would start their engines, while finalizing the boom connection.  During this phase, the upper stage would be following its own trajectory, and the tanker stage trying to match.
7. The booms have built in propellant transfer hoses, and a quick disconnect possibly like the one ULA proposed for cryo prop transfer based on their slip-joint duct design.  The QD would engage, and as soon as it is sufficiently engaged, propellant would start flowing between the vehicles, either pump-fed or using differential pressure.
8. At some point the upper stage gets tanked all the way up.  If there’s still propellant to be transferred, the stages may stick together for a short period of time, with the transfer pump operating throttled back in a way so that the upper stage is not using any of its own propellant.
9. Before the vehicles separate, the QD is disengaged, and the booms reeled back out a bit.
10. Vehicles separate, upper stage continues on to orbit.  Lower stage is still within a glide-based RTLS maneuver of the initial starting site, if the propellant transfer operation can be kept quick enough.

The nice thing about such a setup is that if you do things right, most worst-case failures result in an aborted mission, not a loss of vehicle.  If one of the TSTO pairs doesn’t ignite when air-dropped, you abort (with the upper stage from that TSTO combo having enough propellant to make it home, and you only have to figure out what to do about the first stage).  If you can’t mate up in time, you abort.  If the QD doesn’t work, you abort.  If you can’t keep the vehicles together exoatmospherically, at worst the boom/hose fails, and you use hydraulic fuses to keep that from becoming a loss of vehicle event.  Now, there are many more things that can cause an abort in this scenario, but many of them are things that should get more reliable with practice.  The nicest thing is that many of them can be practiced with first-gen suborbital RLVs without even requiring an air-launch.

Performance and other Observations
Here are some observations from my super low-fi simulations:

1. Performance-wise, this system behaves like a quasi-3STO, giving you a bit of a benefit over a pure TSTO.  For instance, in order to put the same sized upper stage into the same velocity and altitude as the upper stage has at the point the tanker leaves it, you would need something like an 82klb stack instead of two 35klb stacks.
2. The quicker you can get propellant flowing, the better.  The longer you have to throttle back to stretch out the refueling phase, the further you have to carry the tanker stage, which impacts performance.  Ideally you’d like to have propellant transfer done within 90-120s of when the first stages separate.
3. You probably want to throttle-back during prop transfer, this makes it so you’re blowing through less propellant during this phase when you still have a lot of dry mass you’re accelerating.  This would come at the cost of either more lofting earlier on, or more gravity losses.
4. Obviously more exotic propellants for the upper stages tend to provide better payloads, but the quasi-3STO benefits decrease.
5. You want as much of the fueling subsystem mass on the tanker side as possible, since it’s the part that doesn’t get hauled all the way to orbit.
6. You have to be able to pump enough propellant through the transfer hose over the course of the transfer to overwhelm the amount of propellant being used by the main propulsion system.   You are pumping this against the tank backpressure.  Both requirements suggest you probably want an actual pump instead of trying to use pressure feed.

At the end of the day, this would give you an orbital capable system, using the existing WK2 that had all stages fully reusable, and could carry at least 2-3 people.

How to Develop Exo-Atmospheric Refueling
There are three major challenge areas for fielding this technology:

1. The formation flying GN&C development
2. The rapid boom rendezvous system
3. The actual propellant transfer interface

The interesting thing is that the second two systems can at least be tested out in the near-term without even reaching space.  For instance, using two VTVL hovering vehicles (say Xombie and Xoie, or Xoie and SuperMod), you could fly the two together, demonstrate the boom connection and even swap some propellants operationally.   These just require a relatively precise hovering mission, the two vehicles don’t even necessarily need to be actively cooperating.  The next level would involve figuring out how to fly the vehicles in close formation at high speeds and altitudes.  Finally once you had those, you’d reintroduce the exoatmospheric boom rendezvous step.  And if it really looks like it can hook up fast enough, you can demo a little fluid transfer up near apogee, where the air density is low/practically-nonexistent, and the velocities are low.  Only once you’ve developed and demonstrated that, do you try and build the full-up system.

The nice thing is that the incremental cost of flight tests for reusable suborbital vehicles should be really cheap compared to building orbital stages and TPS and other things, so once you have those systems available, it only makes sense to try it (especially if you can find someone crazy enough to pay you to try).

Other Benefits
There are a few other benefits to trying something crazy like this:

• Once you’ve demonstrated exoatmospheric fuel and LOX transfer, people won’t be able to honestly question if orbital propellant transfer is feasible.  This is the acid test.
• The rapid rendezvous techniques end up being very similar to what is needed for a suborbital RLV + MXER tether concept, so by developing this, you open the door for the latter.
• If demand picks up enough eventually to allow you to do a bigger carrier aircraft and bigger system so you don’t need exoatmospheric refueling, having the technique in-hand allows you to launch even bigger payloads in a pinch without having to develop a bigger system.

Parting Thoughts
Complexity should generally be avoided, but sometimes complexity can make things easier, not harder.  A comparable ground-launched TSTO system would likely weigh upwards of 300-500klb wet, and would likely cost more to develop.  It would be a lot simpler, and a lot operationally easier, but its not clear that it would actually be more cost effective or profitable.  Mid-air refueling is still mostly used by the military, but it’s an indispensable part of military operations today, even though it adds complexity.  I think the case of exoatmospheric suborbital refueling will likewise be one of those crazy things that we wonder how we ever lived without.

While fleshing out the technical discussion of these approaches, I also got a lot of useful feedback from commenters and some experts in the industry. In my second post, one of the commenters (a good friend of mine, John Hare) pointed out a fifth RLV approach that I had completely overlooked–airlaunched glideforward TSTO. I plan on making that concept the subject of my last post in the series on RLV approaches–even though there may be other approaches worth discussing, I think the discussion on these five approaches should suffice for now.

The First Stage Recovery Problem: A Review
As we discussed in Part II, one of the key challenges of TSTO RLVs is finding a way to get the first stage back to your launch site without requiring a lot of extra people and infrastructure. It’s also important to get the stage back in a manner that facilitates rapid reprocessing so you can turn the thing around as quickly as possible. While it’s totally possible to do a down-range recovery first stage, and while that may actually be the preferred mode for ELVs trying to transition over into the world of partial reusability/refurbishability, it definitely adds a lot of additional overhead to the operations (more people for running the down-range recovery and return operations), and slows down your possible op-tempo (since recovery and return now take a significant amount of time), as well as placing big restrictions on where you can launch from and to.

In Parts III-V, we discussed the three TSTO options that I had mentioned in my UND lecture. First we talked about pop-up stages, where the first stage intentionally never gains much if any horizontal velocity, so it tends to land back where it came from. Then we talked about techniques that gave horizontal velocity to the upper stage as well as altitude and dealing with gravity losses. We talked about techniques that either used aerodynamic gliding, or rocket propulsion to turn the first stage around and return it to the landing site. The problem is that after a certain amount of downrange velocity (about Mach 3), you end up having to spend a lot of propellant on getting the first stage back. This is propellant that can’t be used for performing the primary mission of putting a payload into LEO. While this mass requirement is worthwhile from the standpoint of easier operations and lower required headcount and facilities, it does come at a respectable cost, which ends up requiring a much bigger vehicle system to place a payload into orbit than a comparable ELV. Ideally, what you would really like to do with an TSTO RLV is to have all of your propulsive efforts going toward putting the payload into orbit, just without the maximum burnout velocity restriction of the glideback TSTO concept we discussed in Part IV. You want the propulsion-free RTLS maneuver without all the restrictions that would normally imply. By launching your TSTO from a carrier aircraft, there is a way to achieve just that.

Air-Launched Glide-Forward TSTO
The concept is actually quite straightforward–John Hare’s suggestion was to launch a two-stage rocket off of a carrier aircraft at a location far enough up-range of the originating spaceport to allow the first stage to reenter after staging, and glide-forward to land back at that same spaceport. This concept manages to meet all three goals discussed earlier: the first stage has RTLS capability, the rocket stack can use performance-optimal staging without the constraints of the glideback concept, and there is no wasted propellant associated with the RTLS maneuver–all of the first stage propellant except for a tiny landing reserve is used for putting the upper stage onto the trajectory it wants.

One of the people who I had read this article before I posted it mentions that launching from the carrier ship at a position uprange of your landing site makes a lot of sense even for air launched SSTO designs. The ability to make it back home in case of an aborted launch is going to be very important for making either SSTO or TSTO air-launched designs work.

Advantages and Drawbacks
While this concept shares some of the advantages and disadvantages of the Air-Launched Assisted SSTO concept, there are a few unique pluses and minuses.

Advantages:

• Obviously, by going with two rocket powered stages, the required performance of the orbiter stage is relaxed a bit, making it much easier to build.
• Depending on how much you need to relax the required performance of the upper stage in order to make the design close, the first stage may actually be in the same performance class as a suborbital booster (such as XCOR’s Lynx). Or, if you go with a more aggressive first stage, it makes the orbital stage mass fraction a lot more achievable.
• Between the performance boost of air launching, and the performance boost of being able to stage along the way, this combination probably ends up with the minimum GTOW (for the rocket stages) of any realistic configuration. While GTOW isn’t everything (dry mass tends to be more important from a cost perspective), it means that you might be able to actually close the case for a manned orbital RLV using a carrier plane as small as WK2. Since the carrying capacity of WK2 is only 35klb (according to a little bird who works on the other side of Sabovich St. in Mojave), as opposed to the 60klb number Wikipedia was providing previously, a manned air-launched SSTO wasn’t going to work.
• By keeping the stages both relatively small for a useful orbital vehicle, fabrication and testing facilities can also be kept small. For a minimal two-person to orbit RLV, you are probably talking about stages not much bigger than SpaceShipTwo–ie something small enough it could be built in a place like Mojave.
• By sizing the stage to fit under existing carrier aircraft (like WK2 or “WK3″ if it ever gets built), you can possibly get away with just leasing the carrier airplane during development, flight testing, and early operations–you don’t have to design your own carrier aircraft or run its development program. You also don’t have to buy the plane outright and staff its operations. Keeping development and early operations costs as low as possible is going to be important.
• Depending on the design of the carrier plane to spaceship interface, it can double as a gantry crane, thus allowing for ground-level maintenance and processing, without needing separate crane systems. This reduces the amount of infrastructure needed to support a launch.

Drawbacks/Challenges:

• Because the first stage has to be able to glide forward to land at the originating spaceport, you are more constrained in where your carrier plane can go for an up-range launch point. This means that you don’t have as much flexibility with regards to orbital phasing for first- or second-orbit rendezvous.
• For near-term carrier planes like WK2, the GTOW limits are probably going to force you to use a cryogenic fuel on the upper stage to make the design close(Methane or Propane at least, but the performance requirements may actually move you towards LH2) . Dealing with cryogenic fuels such as LH2 on the way up to the launch point is going to be tricky, and is going to take a lot of work to make properly. Even keeping the LOX cold on the way up to the launch point is going to take non-trivial insulation systems.
• With a LOX/LH2 upper stage, the rocket might end up being a lot larger than SS2, which might make fitting it onto something like WK2 tricky–this would have to be looked into.
• Your system now has two staging events, one from the carrier airplane and one between the two rocket stages.
• You still need to develop orbital reentry-capable TPS, and depending on the staging velocity of the first stage, it may also require TPS.
• If you’re trying to build to work with something like WK2, you may be down in the size where minimum gage issues start biting you. There are lots of systems that don’t scale down very well past a certain point, and especially on the upper stage portion, those can make life challenging. While staging does reduce the mass ratio requirements on the upper stage, the benefit may end up being largely offset by the fact that you’re shoving the upper stage further into the minimum gage range of the design space.
• Also, even though going two-stage can reduce the performance burdens on both stages by quite a bit, they may still end up being really aggressive mass-wise, particularly on the lower end of the scale.
• Since there are two stages to deal with, the launch licensing process is going to be more challenging, especially since there are a lot more trajectories that would need to be analyzed for E-sub-c, and more potential failure points along those trajectories. Also, since you are more constrained on your launch point location relative to the originating spaceport, you can’t get as far out of civilization before launching (and you have to keep “civilization” out from under the instantaneous impact point as much as possible).

There are probably other benefits and drawbacks, but unlike Part I, I didn’t have the brains of much smarter people to pick while writing this. So this will have to suffice for now.

Thoughts on Various Instantiations and Trades
As I’ve played around with various concepts within this specific approach, I’ve been mostly focused on first-generation systems–i.e. the stuff that could be built in the near to foreseeable future. As you may have noticed, for instance, I’ve been fairly heavily focused on the concept of using WK2 as a carrier plane, especially now that I have some better numbers on its carrying capacity. While it would be possible to design a larger custom-built aircraft that would therefore free-up more mass and make the design easier to close, that would open up a whole can of worms in other areas. It would add another vehicle that needs to be designed and qualified. It would have to go through FAA’s certification process. And it wouldn’t be used as much as a multi-purpose aircraft like WK2 that at least right now has a very large anchor tenant that plans to buy and operate several of them. Which means that the operational burden of owning the carrier airplane may be required right from the start, thus driving up the cost to get this idea off the ground. So for now, I’ve been assuming that WK2 would be the carrier plane, and that therefore the GTOW has to be kept down to 35klb. While it may be possible over time to increase the GTOW capability of WK2, by up-engining the design, I doubt you’d be able to stretch that capacity to more than 50klb. And for now, such upgrades are far off in the future.

Also, I’ve been assuming a payload of approximately 1000lb, which should be enough for a pilot and a passenger, or some mix of cargo. While 1000lb for two people may not sound like a lot, remember that the upper stage already has its own propulsion, GN&C, TPS, RCS, communications, batteries, and other systems that take up most of the weight of a capsule-type system. Mostly for the two-man configuration, you’d be adding in a pressurized cabin, two seats, some controls, some short/medium duration life support, the people themselves, and probably some sort of Intra Vehicular Activity suit (and possibly a hatch of some sort), oh and two parachutes for emergency bailout at lower altitudes/speeds. When two people aren’t needed, this cockpit might be swappable for a small cargo canister, or for some extra propellants for going to a depot. But any way you slice it, the goal would be about 1000lb.

One other common factor in all the approaches I’ve looked at is going with a cryogenic fueled upper stage and a LOX/HC first stage. While there is a lot to be said for avoiding LH2 for an air launched system, its performance advantage definitely makes it look tasty. Plus, with LOX/LH2, you have the RL-10, which is a reliable, high-performance, and readily available engine that is in just about the right thrust class for the job. While a larger carrier plane or in-air refueling might allow for an all-LOX/HC design, both of those add complexities that I would like to avoid if possible. You might be able to make an all LOX/HC vehicle in the size range we’re talking about work if you had a really high performance engine, but you’d still be running into minimum size issues with a lot of components. While it is harder to achieve a given propulsion system mass ratio with LOX/LH2, you have more mass to work with, and a fully-functional orbital stage requires a lot of other subsystems that don’t scale down linearly. For instance, on the Centaur stage, the tanks and engines only weigh about 40% of the stage mass. The rest of it goes into several other subsystems (batteries, electronics, RCS, structures, etc) that while scalable are not as directly tied to propellant density as tanks and engines are. Using the estimates I got from some LM friends, the actual LOX/LH2 tanks and engines would take up less than 1200lb for most of these concepts, leaving quite a bit of mass for those other subsystems. An all LOX/HC approach *might* work, but you’d only save about 500lb on the propulsion structure, and you would have about half as much dry mass to play with for the rest of the subsystems (and that’s assuming 350s Isp on the upper stage, which is pretty impressive as far as LOX/HC engines go). It isn’t clear to me that LOX/LH2 for the upper stage isn’t actually the easiest way to go.

Now, if Scaled is able to increase the performance of WK2 any (or if it turns out that the carrying weight constraints can be relaxed if necessary), that may change things, but for now, with the limited GLOW available, I think LOX/LH2 for the upper stage is probably easier (if you can solve the thermal issues).

Based on those constraints (35klb gross, 1klb payload, LOX/LH2 single RL-10 upper stage, LOX/Hydrocarbon FS), I’ve looked at a few approaches.

The most straightforward would be a serially staged system, where the first stage fires, burns to near depletion (leaving a little propellant for landing), the two stages separate, then the upper stage lights its engine and continues on to orbit. For this system, you want a first stage with a decent amount of thrust. Previous studies on air launched SSTOs talked about a T/W of ~1.4 being desirable. That would equate to about 50klb. That’s bigger than the C&Space Chase-10 engine, or the Air Launch QuickReach upper stage engine, but smaller than the Merlin-1C. Depending on how aggressive you want to go with the first stage design (and whether you want a pilot on that stage or not), you have a range of options. On the conservative, piloted side, I ran some BOTE calculations that suggested you could possibly have a first stage that would have about the same dry mass (just a bit less due to being a 1-seater, and being a little more aggressive on the materials and engine T/W ratio) and propellant loading as a XCOR Lynx Mk1.

This would result in an upper stage that had about 42% the propellant loading of a Centaur upper stage, while having about 78% of the dry mass, not counting the payload. This gives a propellant mass fraction of just a hair over 80%, which while aggressive for a reusable stage isn’t off in fantasy land. The burnout velocity for the first stage would actually be under Mach 3, which means you could actually launch right over your launch site and glide back, or operate anywhere within about a 50miles diameter. By having your launch point be so close to your orginating spaceport, it might make it easier to get launch licenses for the various trajectories you would want to fly.

OTOH, by going with a more aggressive first stage (say LOX/Methane, no pilot, more aggressive materials, higher tankage fraction, etc) you can reduce the required upper stage delta-V by quite a bit, which means that with the same 3500lb of dry mass to work with, you might be able to get the upper stage propellant load down to about 31% of a Centaur loading (about 14klb), which is a more manageable 75% pmf. This would have a much higher burnout velocity (above Mach 5), which would imply needing to launch several hundred miles uprange. Which way you go depends on the tradeoff between easier development and operations of the first stage, and the challenge of getting the upper stage to design to close.

[Just for kicks and grins, and because I know that most alt.spacers will decry the mere suggestion of using hydrogen, here's what you would need to get a LOX/HC two stage combo to work. You'd need high Isp engines: 340s Isp for the first stage and 350s for the upper stage. And you would need dry masses of about 2000lb for the first stage and 1900lb for the second stage--not counting the 1000lb payload--and propellant masses of about 15000lb each. In other words, the first stage would be very similar to the first stage shown above, but the upper stage would also have to be very aggressive. That's about an 87% pmf. It's not impossible for vehicles this size, but I think it may end up being a lot more challenging than the LOX/LH2 design.]

The other main branch of the tradespace I’ve looked at involves using just a single main propulsion engine (the RL-10) modified with TAN injection and using propellant crossfeed from the first stage. You could either use the cross feed just the TAN propellants, or the TAN propellants and the core oxidizer. Both have benefits and drawbacks. Neither case ends up getting you results drastically different from the serial staging route, but you only have one main propulsion system to deal with. If you cross-feed the RL-10′s core LOX during the burn, you can actually reduce the required size of the upper stage a bit, but at the expense of shifting the stage mixture ratio more towards hydrogen, which will cost you density-wise. The benefits compared to a serially staged system are fairly modest compared to the added complexity, so it might not be worth the hassle.

One more thought on instantiations. There’s nothing per se that requires an air-launched TSTO system to be a winged HTHL configuration. In fact, it might be possible to do it with two VTVL stages. One air launch paper I saw on L2 of NASASpaceflight.com talked about a massive air launcher for a Delta-IV type launch vehicle. They found that if you light the rocket engine with the airplane still attached, that you can get enough thrust to pitch the aircraft up to a high-enough angle before staging that you eliminate the need for having wings on your stage. It costs you a little extra propellant for a few seconds, but if it ends up making it easier to achieve your mass ratio, because you don’t need wings, all the better. Your stages wouldn’t have as much cross range, so your launch locations would be a little more constrained, but this is still probably doable. Also, for VTVL stages you would either need landing engines, or some sort of altitude compensation or flow-separation control in order to allow you to use high-expansion-ratio engines for the main acceleration for landing. Just food for thought.

Related Technologies
Like the other RLV approaches, there are several related technologies, and areas where we need to gain more experience with before such a design can close. Here are what I see as the key related technologies, and a bit about how I would go about filling in the knowledge gaps:

1. TPS–This shouldn’t come as a surprise to you. Robust, reusable, survivable, and lightweight TPS is pretty much the key to any orbital RLV. As mentioned in my previous articles, there are lots of ideas out there, but precious little data on all but a few of them. Using a platform like WK2 coupled with some sort of kick stage, it should be possible to test out various TPS approaches, gradually expanding the envelope. The process for systematically testing out orbital TPS in a low-cost, rapid-iteration manner is probably a topic with more discussion at a future date.
2. Cryogenic Propellant Storage–Even if you don’t go with LH2 for the upper stage, you’re still probably stuck with at least one cryogenic propellant in order to get the performance you need. The long trip from the launch site to the uprange launch point is going to expose your vehicle to a lot of convective heat transfer on the outside. Designing proper passive and active cooling systems that are lightweight enough to carry to orbit (or which can be located on the carrier plane with the minimal amount of interconnects) is a tricky problem. Fortunately, in many ways its related to other cryogenic fluid management challenges that need to be solved for on-orbit applications. Low conductivity tank-to-frame connections, internal insulation, and possibly various active cooling techniques will all need development. Fortunately, you don’t really need even an honest-to-goodness rocket stage to test these technologies. It might be possible to make a testbed that’s just a big propellant tank that you could fly using WK2 to gather data before you jump into the full development program. By being able to retire the risk and mature the design with a simplified non-flightweight prototype, you can get answers to your key questions a lot quicker. Even with all these precautions, some sort of onboard tankage and propellant conditioning equipment will probably be needed, but the optimal solution is likely going to be a mix of passive vehicle-side stuff combined with active mothership-side hardware.
3. Airframe–At least for the upper stage (and also for the lower stage in most of these configurations), it is going to be quite a challenge packing that much hardware into that small of a mass budget. This isn’t as much of a problem for VTVL vehicles, which tend to have much simpler structures that use the tanks themselves for much of their strength. Fortunately projects like XCOR’s Lynx MkI and MkII will help provide some experience on lightweight, high-propellant mass fraction winged vehicles. Sure, they’re only suborbital, but that’s the point. It’s a lot easier to take a reliable, working system, and then start finding ways to make it lighter, than to try to design a super lightweight system from scratch and then make it reliable. A suborbital winged vehicle will require most of the subsystems that a winged orbital stage will require, but it can serve as an incremental step along the way–showing where the high-mass components are, and allowing you to start figuring out what it will take to close the design. More importantly, it may turn out that the data from such an excercise may show that an orbital vehicle this small just ain’t going to happen. That’s useful information too, even if discouraging. Much better to learn if something looks feasible from a smaller, profitable project, than to toss a bunch of money into something you don’t know will work.
4. Orbital Prox-ops Tugs–For any small RLV, you’re usually fighting pretty hard agains minimum gauge issues and such. Anything that can allow you to remove hardware from your system, and transfer it to something that can be left on orbit, is a huge win. It might be possible to allow the prox-ops tug to carry almost all of the mass associated with rendezvous, docking, and even people transfer. Imagine a tug with multiple arms and a “transfer tunnel”. The tug goes out, rendezvous with RLV, grapples it with one or two handholds, and brings it back to the station. Say its transfer tunnels have extendable portions (kind of like the loading ramps for commercial jetliners), that has an electromagnetic ring with a small ferrofluid reservoir. You have a ferromagnetic matching ring built into your vehicle around the crew ingress/egress hatch. The transfer tunnel seals against that ring providing a vacuum tight seal, using the robot arms to articulate things correctly. The tug then hauls you up to the station, grabs a handhold or two located near an EVA hatch, and repeats the process. The people could then leave their vehicle wearing their IVA suits. They’d enter through the airlock itself. That way if there’s any problem with the sealing, it doesn’t endanger the station, and the vehicle is designed to operate even with a loss of atmosphere…Just sayin.

   Simpler tugs, possibly without the transfer hatch could also be of great use for other types of missions (propellant transfers, delivery of external spares, delivery of small satellite kick stages, etc). One of the keys to propellant depot operations is going to be minimizing the amount of hardware necessary on the tanker vehicle. If each tanker vehicle has to be built as its own Autonomous Rendezvous and Docking robot, with full RCS suite, full GN&C suite, etc, there’s no way you’re going to get costs down or propellant efficiency (ie propellant mass as a percentage of the orbited payload mass) to a reasonable level. What you really want to do is have the propellant tanker literally just be two big tanks with a little plumbing and some grapples. The stage that launches it provides attitude control long enough for the tug to come and snatch the tank delivery and haul it up to the depot. Then, you want the plumbing on the tanker side to be the minimal stuff to store and handle the fluid between launch and delivery to the depot. You want as much of the transfer related hardware as possible on the tug or station side. Ideally, I’d have the plumbing on the tanker side just be a panel with some grapple points, a couple of manual lever actuated ball valves, and some quick disconnect receptacle fittings, and have all of the smarts and all of the robotic bits on the part that already has to have articulating arms and end effectors. Make the whole thing so simple and cheap that the State Department will give you ITAR permission to sell it to anyone in the world to integrate into their own dumb tanker module.

   But the key is that by not requiring all of the specialized hardware normally associated with rendezvous, docking, and fuel transfer to be hauled around on the vehicle for every flight, it makes it a lot more likely that you can close the design case. The only drawback is that you really need someone to develop a good tug at that point, and unless they can find an existing market to justify their tug’s existence before you start your RLV project, you might end up in the unenviable position of needing to develop both the RLV and the tug at the same time–thus greatly increasing your odds of not being able to pull it off.

5. Thrust Augmented Nozzles–In situations like this where you need good Isp, but you also need good propulsion T/W ratio (to make the mass budget close), having a technology like TAN that helps you on both scores could be very useful. Also, TAN as mentioned earlier, can enable you to do a project using only one main propulsion engine, provided you can cross-feed propellants reliably. Right now, TAN has been proven out on a small, laboratory scale, but there’s still work to do to bring it to primetime.
6. Composite Cryogenic Tanks–At least for the LOX, the ability to store some of the propellants in the wings allows you to save some weight for the vehicle (since now you’re combining two structures into one), and also probably helps with controling your vehicle’s CG. Cryo composite tanks are also likely going to have integral insulation, and thus be much better at reducing boiloff. I don’t think XCOR’s “Non-burnite” is rated down to LH2 temperatures, but even if only your LOX tank can use the technology, it’s still a big win. And for LOX/CH4 or LOX/subcooled-Propane designs, you can use it for both tanks.

The Path Forward
Like all of the other proposed RLV approaches, the path to orbit lies in suborbital vehicles. Regardless of stupid arguments about differing energy levels required and all that crap, the reality is that suborbital RLVs end up having to develop almost all of the subsystems you need for orbital launch vehicles. And you have to integrate them into a working, reliable package. And you need to do so in a way that can have quick turn-around operations. Sure, orbital operations are going to require lighter airframes that have more propellant in them. And more powerful and efficient engines, but many of the lessons still apply. The interesting thing is that on the scale we’re talking about for this specific approach, both stages end up being not much bigger than the XA-1.0/1.5 concept we’re working on at Masten or the Lynx concepts working on at XCOR, and in fact the first stage for such a system, as I’ve shown earlier, may very well be on the same performance level as either of our or their vehicles. So, a lot of that work is just continuing down the path that we and several other companise are already going.

The other big path forward would be using the WK2 platform as an engineering testbed. Both for testing out cryogenic storage and Airborne Conditioning Equipment for keeping those tanks toped-off till launch, and also for launching small stages carrying TPS experiments. Though honestly, many of the suborbital companies in development are also looking at converting their vehicles into nanosat launchers, which would be just fine for launching TPS test hardware. For airlaunched VTVL style vehicles, work could also be done to demonstrate both the ground handling aspects (tipping the vehicles over so they can be mounted to the mothership), and the launch aspects (lighting the VTVL stage, doing the pitchup maneuver, and separation). Lastly, they could always work on ways to try and increase the cargo capacity of WK2 a bit, to make our lives easier. Maybe using the airlaunched first stage as a JATO bottle?

One last big challenging area of development is going to be working with the AST to get such a system licensed in a manner that allows it to operate as flexibly as it needs to. One suggestion I heard was that if this technology were being developed for the government first, that it could build-up reliability experience to the point that the AST would be more willing to work with it. The other option is that the data built up from WK2/SS2, Lynx, and other suborbital vehicles can also be used to make the regulatory case easier. Right now, to be conservative, the E-sub-c calculations are done assuming that your vehicle will crash and die every single time you fly it. But once you’ve flown something several times, you’ve demonstrated a certain reliability level, that can be factored into E-sub-c calulations, thus allowing you to operate over slightly more populated areas. If you’re launching uprange of your landing site, and gliding forward to landing, odds are your Instantaneous Impact Point at separation is going to be fairly close to your launch site. If you design your trajectory so that staging occurs with an IIP slightly past the originating spaceport, then most of the challenge on the launch licensing is going to involve your first stage operations, as most orbital spaceports will probably be sited such that there isn’t much population nearby to the northeast, east, or southeast. For the first stage operations, since it really isn’t much more than a suborbital rocket, getting lots of safe flight experience with such a combination, over unpopulated areas should be fairly easy to do. Once that’s worked out, it should be much easier to make the E-sub-c numbers look reasonable enough that the AST will allow you a loose enough launch license that you can move your uprange launch point around in the way you’d want to to really take advantage of air-launch.

Conclusions
As I mentioned at the start of this post, this is my last post for now in this series. My main goal was trying to show people that there are multiple realistic ways of solving the reusable earth-to-orbit transportation. I also wanted to introduce people to some of the challenges and tradeoffs inherent in developing such systems, and show try to paint some sort of a vision of the path forward from here. Lastly, I wanted to lay a sort of framework for discussing the business aspects of orbital access. I hope you all enjoyed this series, and I think everyone for their comments. Now its time to roll up our sleeves and get back to work.

Two Forms of Boostback Techniques
In a Boostback TSTO system, the first stage provides not only vertical velocity to overcome most if not all gravity and drag losses and significant downrange velocity, but it also provides enough propulsive capacity to return itself to the launch site after separation. Unlike the glideback case, the Boostback TSTO approach stages at a sufficiently high velocity that at least some of the return to launch site (RTLS) delta-V has to be provided propulsively by the stage itself. Also, unlike the glideback approach, the first stage does not have to have a high L/D ratio, and in fact boostback can be used with VTVL vehicles.

The first, and most well-known form of boostback, (the form proposed for use with the Kistler K-1 vehicle, which I’ll call Propulsive Boostback) involves a first-stage rotation maneuver after staging, followed by firing the engines long enough to both cancel out all of the downrange horizontal velocity, and provide enough net uprange horizontal velocity that the stage can land back at the launch site. In the case illustrated in the presentation I linked to in the previous part (and further detailed in this report), the optimal staging velocity was found to be about Mach 5.2 (~1800m/s), at an altitude of around 52.5km, and a staging flight path angle of about 31 degrees. For this case, I did a little analysis, and I’m estimating that between the ascent phase and RTLS boostback maneuver, the total first stage delta-V would be around 5500-5800m/s. But the good news is that the upper stage would also be down in that range (ie slightly lower than 6km/s even including landing propellant for the VTVL case). The Kistler K-1 vehicle used a similar but slightly different trajectory, where the staging was planned to take place at about Mach 4.4 (~1500m/s), and at around 42km. That would result in a slightly higher required upper stage delta-V requirement, but a lower first stage performance requirement. This figure, from Barry Hellman’s report I linked to above shows an example propulsive boostback (starting at the staging point):

While Propulsive Boostback is the most well-known form of Boostback, I realized last week that there was another approach that is also uses a form of boostback maneuver. For sake of clarity, and for lack of a better term, I’ll call this approach Lift Assisted Boostback.

I thought of this boostback approach in response to some questions to my previous post on glideback approaches. Someone had asked why you couldn’t stage at an even higher velocity. I started in on an explanation about how at velocities any higher than Mach 3.2 (using the assumptions in the prior studies), the rocket would not be able to glide back all the way to the landing site, and that therefore you’d need some sort of additional propulsion event after staging in order to get home. While people typically recommend turbojets for such missions (thus switching from glideback to “flyback” for the first stage), I suggested that it might be worth just using the rocket engines in such a situation. Upon further thought, I realized that there might be more to this suggestion than I had originally thought.

Basically, if the first stage has a sufficiently good L/D, what you can do after staging is, glide downrange a bit, and then perform a turn-around maneuver aerodynamically (once you’re back in the atmosphere enough to do so), and finally relight the engines to provide enough momentum to get you back within glide back range of your launch site. By performing the turnaround maneuver, you’re using aerodynamic lift to bend your trajectory around so that the downrange (away from the launch site) velocity is now actually turned into velocity heading back home. That way, when you light your engines for the boostback maneuver, while you may be at a lower altitude, you no longer have to null-out the downrange velocity, and your propulsion system also doesn’t have to provide all the uprange velocity in order to return to the launch site.

[Update 7/1/08: A commenter mentioned that there's a third approach that combines some of the features of propulsive and lift assisted boostback to avoid some of the key drawbacks of both. Basically, if you have a vehicle that both has good L/D, and has a propulsion system that can handle a boostback retrofiring maneuver, you have a third option that avoids hypersonic flight and excessive TPS requirements, while also keeping the first stage Delta-V more reasonable. Basically, after staging you immediately pitch over and decelerate until you've slowed yourself down enough that you can reorient yourself and do a glideback trajectory from there. While it adds some extra operational complexity (two rotational maneuvers), it gets rid of the TPS issues with lift assisted boostback, and gets the required delta-V for the stage down into the 3.8-4km/s range instead of the 5.6-6km/s range required for a purely propulsive boostback technique. Food for thought.]

Benefits and Drawbacks of Propulsive Boostback
The two different boostback techniques have somewhat different advantages and drawbacks. Propulsive Boostback is the form best known, so I’ll discuss some of the pros and cons of this approach first.

Benefits:

1. A common benefit of both approaches over the previously discussed methodologies is that the delta-V requirements on the upper stage are much lower. Depending on the exact staging conditions, the upper stage may need to provide as little as 5800m/s, compared with at best 6400m/s for Glideback TSTO, 7400m/s for Pop-up TSTO, and 8000m/s for Air-launched SSTO. 5800m/s equates out to a propellant mass fraction of about 0.83 for a medium-end LOX/Kerosene stage, and about 0.73 for LOX/LH2. Both of these are very realistically attainable pmf values.
2. The delta-V requirements put the two stages at a level of technology only slightly beyond that needed for small suborbital vehicles (which tend to suffer from higher drag losses than larger suborbital vehicles, and thus need a higher total delta-V for the same apogee), making the step from suborbital to this form of orbital easier.
3. A boostback TSTO has the option of doing occasional downrange landings (if there is a suitable landing site) in instances where you need to lift heavier payloads.
4. With the upper stage empty an unfueled, the first stage could actually self-ferry the stack fairly long distances (several hundred miles).
5. The boostback maneuver ends up resulting in a very low reentry velocity compared to what you would expect from the staging horizontal velocity. The reentry velocities are low enough, ~Mach 2, that TPS is almost unneeded for the first stage.

Drawbacks:

1. The first stage ends up requiring a lot more delta-V than earlier methods, but a substantial chunk of that is used for the RTLS maneuver. At low achievable propellant mass fractions and Isp, this results in a much easier to build RLV than the other approaches. However, as the achievable mass fraction and Isp increases, at some point the extra delta-V actually makes the vehicle heavier (both in total mass as well as in just dry mass) than a pop-up or glideback stage. While admittedly higher dry mass doesn’t necessarily equate to higher costs (a 1000lb dry mass stage made of 5383 TIG-welded aluminum is going to cost a lot less than even a 500lb dry mass stage made of friction stir welded Li-Al alloy, or a 250lb stage made of Unobtanium Wishalloy-X), there may be a performance point at which the boostback design no longer has sufficient cost or performance advantages over glideback or pop-up designs to justify the more complicated maneuvers.
2. The turnaround and boostback maneuvers (often called the RTLS maneuvers) are somewhat complicated, and involve in-air relights of engines. Admittedly for a VTVL stage, your propulsion system better be rock-solid reliable anyway, so this isn’t as big of a deal for VTVL boostback systems, but every additional complication comes at a price.
3. Boostback trajectories have more of their safety-critical operations occurring downrange of the launch site than many other approaches. This means that more attention will have to be paid during launch license applications to making sure the trajectory is tuned to keep the risk to the uninvolved public low enough.
4. More to the point, at some point, the Vacuum IIP (the point where the vehicle would hit if it’s propulsion systems failed at that instant and there was no atmosphere) ends up loitering over some downrange site. Making sure you can have this occur over an unpopulated area is critical for getting launch licenses.
5. Trajectory tuning like this requires extra performance margin. With enough margin, you can probably find appropriate trajectories for most launch sites and azimuths, but the more generally useable the stage wants to be, the more margin you need. The problem is that the first stage in this case is already getting near the steep part of the delta-V vs. Mass Ratio curve. Adding extra margin becomes harder and harder very rapidly.

There are probably other benefits and drawbacks I’m not thinking of, but these are a start.

Benefits and Drawbacks of Lift Assisted Boostback
While there are several big potential advantages to the Lift-Assisted Boostback, there are also some unique differences and drawbacks. Unfortunately, since this isn’t a concept I’ve seen investigated in the literature before, and as the aerodynamic turn-around maneuver is more complicated than I know how to easily analyze (and I don’t have access to a full-up 6DOF trajectory analysis program), I will only be able to give some general thoughts. If anyone reading this actually has enough time to analyze the concept in detail, they might be able to provide some more insights.

Benefits:

1. By using aerodynamic lift to do the turn-around maneuver, you will end up requiring less RTLS delta-V for a given staging velocity.
2. While it is possible to do a propulsive boostback with an HTHL stage, all of the main burns for a lift-assisted boostback system are performed at altitudes where aerodynamic control surfaces can provide some or all of the control, thus allowing you to use engines as simple as those that would be required for glideback.
3. This approach gives you most if not all of the reduced upper stage delta-V requirements that a propulsive boostback technique without anywhere near as much of a first stage delta-V penalty. This means that this approach may stay competitive with glideback and pop-up approaches even as the level of achievable stage performance increases.
4. Unlike propulsive boostback, your IIP never ends up stopping and loitering over any given point, because your trajectory is being bent around aerodynamically. A rapidly-moving IIP crosses a given chunk of land faster, thus making it easier to maintain a reasonable E-sub-c for launch license purposes.
5. The fact that this approach doesn’t really require any unique capabilities not needed for glideback (glideback may assume that you have the capability to relight the engines in case you need to do a go-around at the landing site), means that you can incrementally upgrade a glideback vehicle to be able to perform a lift-assisted boostback. For a given glideback TSTO design, as you incrementally add first-stage performance, that offloads performance requirements from the upper stage, allowing it to carry more payload over time.
6. Most of the aerodynamic maneuvering occurs at a high enough altitude and speed that it’s possibly in the hypersonic regime. In the hypersonic regime, lifting bodies are just about as good as winged stages, which means it might be possible to have a VTVL system that has a lifting body configuration. You’d use the lift for aiding in the turn-around maneuver, and some of the glideback, but would use propulsion for takeoff and landing. Thus getting some of the benefits of a winged vehicle, while avoiding the disadvantages of a VTHL system.

Drawbacks:

1. In order to do the turn-around maneuver, your stage is going to be going fairly fast during reentry, and in order to maximize performance, you will likely end up exposing your vehicle to pretty ugly thermal environments–much worse than propulsive boostback, glideback, or pop-up TSTO designs. Nowhere near as bad as orbit, but possibly as bad as “flyback” trajectories. This requires a real, honest-to-goodness TPS system that will need to be developed and proved out. We’re talking maneuvers going on at airspeeds faster than the SR-71, so this isn’t a trivial problem, even if the duration is relatively brief.
2. Unlike propulsive boostback, if you staged at a similar velocity, you’d end up going much further downrange before you could get back into the atmosphere far enough to start turning around. Depending on how much of the velocity you can maintain after the turn, this may require a significant burn to get the vehicle back to the launch site. In other words, at least some of the benefit you get from not having to use propulsion to null-out the forward velocity is counterbalanced by possiblly requiring a bigger burn to get up to speed to get back to the launch site. This may mean that the optimal staging point is at a lower velocity than for propulsive boostback. Or it may just mean you have to do a hotter turn-around maneuver.
3. Since you end up going much further downrange, it may be harder to find areas remote enough to launch out of.
4. A failed engine relight may force an emergency landing a long way from your launch site. This may require a decent amount of contingency planning.
5. Doing a large hypersonic turnaround maneuver may end up causing a large sonic boom, which may also complicate trajectory planning.

There may be some other benefits and other problems, but those are the major ones.

Enabling Technologies and The Path Forward
Boostback TSTO designs share similar enabling technologies to the other approaches. HTHL versions could really use composite cryo tanks to allow them to fly with “wet wings”. All of the different boostback approaches can benefit from suborbital vehicles–it may even be possible to test out a lot of the techniques necessary using suborbital vehicles. The orbital stages for these approaches need TPS work just as much as any of the others–but in the case of lift-assisted boostback, even the first stage will require advanced TPS work. Altitude compensating nozzles (or Thrust Augmented Nozzles, which also have a form of altitude compensating) help a lot, as most of the RTLS burn is done at high altitudes, and for propulsive boostback, higher thrust for the boostback maneuver ends up reducing the required delta-V back by a small but not insignificant amount.

The real way ahead for both of these projects is going to involve testing out the required maneuvers with suborbital vehicles first. There are some groups in the Air Force that are really keen on using this technique as well, and they have been pushing it quite hard lately. Even sub-suborbital vehicles (like XCORs Lynx, most of MSS’s XA-0.x demonstrators after 0.2, and most of Armadillos’ nearterm vehicles) can do some of these experiments, and it would be good if the Air Force could continue working with these firms as their vehicles become available. Admittedly, I’m somewhat biased there–being a propulsion engineer for one of the companies that could benefit from such a move. But by using a boostback maneuver with a suborbital sized vehicle, the delta-V requirements for an expendable upper stage would be low enough to allow for a decent nanosat launcher (or a vehicle that could launch TPS testing reentry vehicles, which would be a great way to get the data you need before you can start building an orbital LV.

So, does anybody have a 6DOF simulator and lots of time on their hands that wants to do some extra analysis of this lift-assisted boostback maneuver? It might make for a fun Master’s Thesis.

This morning, while I thought back to the UND lecture that started this all, I realized that the key goal of this series has always been to show that there are several realistic approaches to doing RLVs, and to try and give a high-level overview of the different approaches and how to get there from here. While there is definitely a lot more detail that I could go into on each of these topics, they’re not the only ones I want to write about, and I just don’t have the time to both go into the level of detail some would prefer while also being able to do much of anything else. If someone is interested in taking what I’ve got here, and fleshing these out into a more formal and detailed form, let me know. Otherwise, I’d like to just continue as I’ve been going with giving a high-level overview of the most promising orbital access techniques I’ve been looking at.

In Part III, we discussed a TSTO approach where the first stage provides only vertical velocity, and the second stage provides all the horizontal velocity. As many have probably notice however, requiring the upper stage to provide all the horizontal velocity makes the upper stage design a lot more challenging, and also tends to drive the overall vehicle size up substantially. The obvious question is, are there ways of having the first stage provide horizontal velocity, while still returning to the launch site? It turns our that there are some ways of doing that, and this post will focus on the first, and by-far easiest of those methods: glideback.

Glideback TSTO: An Introduction
As has been mentioned several times in this series, the rocket equation is an exponential function. As you near the “right-side of the curve”, ie higher velocities, the engineering challenge of building a reusable stage becomes rapidly more difficult. The corollary of this is that by moving the velocity requirement for a stage even slightly lower, the gains can be quite large. For instance, by going with air-launch, I showed that making a functioning “assisted SSTO” may actually be achievable with near-term available technologies, while a ground launch SSTO is still a much harder challenge. Likewise, even though the pop-up TSTO approach only saves the upper stage about 600m/s over the air-launched SSTO approach, it too makes a big difference. So, at least on the “part of the curve” (delta-V versus mass ratio) that we’re looking at for an orbital stage, adding even a small amount of downrange velocity can still have a very large, and beneficial impact. The challenge is doing so while still maintaining the operational advantage of being able to have the first stage return directly to the launch site at the end of its mission.

The easiest way to accomplish this is by using aerodynamic lift. The idea behind glideback is that the first stage takes the upper stage up to a certain altitude, and gives it some downrange velocity, then it stages, decelerates a bit, turns around, and glides back to the landing site. Naturally, the better the L/D of your system, the more delta-V the first stage can impart while still making it back home, so while it may be feasible for a VTVL first stage to take some advantage of this technique, it is more naturally suited to HTHL approaches.

This presentation, done by Barry Hellman of Georgia Tech, provides some more details on this approach (as well as the boostback approach which will be discussed in Part V), and more details can be found by googling “glideback” or by searching for “glideback” on NASA’s NTRS site. The idea was previously investigated as part of the Shuttle II or Future Space Transportation System studies done in the late 80s and early 90s. The basic concept is that the two stages take-off horizontally, accelerating to about Mach 3-3.2 (~1100m/s) at an angle of about 45 degrees (thus providing about 775m/s of horizontal delta-V), and then staging at an altitude of about 32.5km. At that point, the first stage performs a high-alpha reentry to slow down a bit, turns around, and glides back to the launch site for an unpowered horizontal landing. Mach 3.2 was chosen as the optimal point for the Shuttle II analyses (though I’m not sure all of the assumptions going into that number), as going much faster would preclude being able to return to the landing site on gliding alone. There are different variations on the theme that are possible, and different assumptions will yield different burnout velocities, angles, and altitudes (ie Your Mileage Will Vary), but that was the basic idea.

Shuttle II Glideback Concept from NASA Technical Paper 3335:
Analysis of the Staging Maneuver and Booster Glideback for a Two-Stage, Winged, Fully Reusable Launch Vehicle

Benefits
So, what are the benefits of this approach compared to the other ones we’ve discussed so far?

1. The first stage in this approach is actually imparting a significant amount of horizontal delta-V (almost 800m/s), thus making the upper stage’s job much easier.
2. This approach takes a lot more advantage of the benefits of HTHL approaches, in that it’s using wings to lower the required takeoff T/W ratio, and using the wings to do a lifting ascent.
3. The engines on the booster stage can be much simpler than for a VTVL booster stage. You might not need throttling or gimballing, thus allowing for a much simpler propulsion system–if MSS had been doing HTHL, and if it had had access to an airframe, our engines were mature enough two years ago that we probably could’ve had our own EZ-Rocket flying for over a year now.
4. Due to the lower delta-V requirements on the upper stage it becomes much easier to make the upper stage use a denser propellant combination, without taking as much of a hit for the choice.
5. The reentry velocity for the first stage is even lower than most suborbital vehicles, thus completely eliminating the need for any first-stage TPS.
6. The first stage doesn’t require a very high mass ratio, thus making it quite low-tech. While much larger than an XCOR Lynx Mk II, the vehicle would only need technology on-par with the Lynx Mk I to be workable–ie the technology risk is very low.
7. HTHL vehicles tend to provide for much more graceful abort modes. For instance, a total propulsion failure of the first stage might not even require stage separation. You might just dump oxidizer, and then glide back to a landing. Fixed engines are much easier to “armor” against hard starts (and are much easier to make more deterministic than a throttling engine, thus making hard starts potentially less likely).
8. Due to the low-Mach number, and low required Mass Ratio, the first stage has much more in common with a normal aircraft than a launch vehicle–it can borrow heavily from aircraft construction techniques and some subsystems, thus leveraging a more highly matured transportation industry.
9. Depending on the flight trajectory taken, the first stage might not actually meet the AST definition of a suborbital rocket. While it isn’t clear why you’d want to have the first stage regulated by the other part of the FAA, if you wanted to, you probably could force the trajectory either way depending on which you thought was more commercially useful.
10. HTHL vehicles like this are more likely to be able to operate out of existing airfields. While operating out of LAX anytime soon is unlikely, there are plenty of large airfields out there that could easily attain the required FAA launch site licenses by leveraging work done by the Oklahoma and Mojave spaceports (not to mention just using Mojave or Oklahoma spaceports). This flexibility makes it easier to operate out of multiple launch sites not necessarily tied to existing (and expensive) launch ranges.
11. The first stage operating by itself without a fully-fueled and loaded upper stage on top probably has enough propulsive power to make several hundred miles downrange. It can also probably do so while operating as a rocket powered aircraft, thus making it easier to self-deliver the first stage to a given destination. Once again, how much the FAA would appreciate someone trying to do this is left as an exercise for the sufficiently masochistic reader.
12. The first stage has a low enough required MR that you can probably include hardware, such as ramps, that would allow an unfueled upper stage to be remounted to the first stage without the use of a crane. Sure, that goes against standard aerospace weight-minimizing practice, but if it allows cheaper and easier operations, it might be worth it. Any time you can allow for ground level servicing, maintenance, and inspection, it makes operations a lot easier.

Once again, there may be other advantages I’m overlooking, but those were some of the key ones that I could think of.

Drawbacks, Limitations, Constraints, and Challenges
As you probably guessed, there are some drawbacks to this approach in general, and the specific implementation mentioned above. Unlike the Pop-up TSTO approach, there’s a bit more flexibility on the exact trajectory, which means that some of these issues may be resolveable by using clever trajectory planning.

1. The staging velocity and altitude result in a fairly severe dynamic pressure environment during stage separation. 800psf to be precise (38.3kPa for our metric-using friends). This makes staging a lot more dicey. The article that I pulled the picture from includes some analyses on how to solve this problem, but it still has a fairly high associated pucker factor. It may very well be worth redoing the analysis with staging dynamic pressure being given a higher weighting factor (ie. at the cost of some performance).
2. Staging at 32.5km at that speed and angle also means that your first stage apogee only reaches a little over 60km. That means that you’re going to take some gravity losses with the upper stage unless the T/W ratio is really high. This is especially the case if you coast up to a higher altitude to do staging. This will slightly reduce the benefit of the downrange velocity. It might be possible to change the trajectory such that the first stage apogee is 100km, and the staging point is over say 50km to keep the dynamic pressure down, while still keeping some or all of the downrange velocity, but I’m not in a good position to say what the tradeoffs would be.
3. The wings and landing gear for the first stage have to be designed to handle lifting the full stack, and for doing emergency landings. Fortunately the first stage doesn’t need very high MR, so this isn’t as big of a problem as it would be for a ground-launched SSTO for instance.
4. If the first stage is running a trajectory that causes it to be classified as a launch vehicle, it’s IIP will stop over some downrange point. Also staging occurs with the IIP at some downrange point as well. It will be important to try and locate this point such that it isn’t over populated areas. This may limit somewhat the available launch azimuths, and may require the first stage to have some extra performance margins in order for different launch locations to shape the trajectories to minimize the E-sub-c for the flight.
5. There are also issues with scalability. While the NASA study mentioned previously was for an HLV sized vehicle, realistically, it’s going to be a challenge getting anywhere near that big anytime soon.
6. The orbital stage TPS problem. Same as with the other approaches, but as the stage gets lower delta-V, it also becomes slightly less fluffy, which tends to increase the TPS material challenge.
7. Glide landings are no fun, but depending on the engine concept, it should be possible to do what XCOR does, and have propellant on-board and the capability to do a “go-around” burn. As it is, it’s been fun over the past few weeks watching a certain “Undisclosed Flying Object” do multiple in-air relights (and some pretty sweet maneuvers) over the Mojave Spaceport. For some reason I think that making the landing not have to be a glide landing wouldn’t be that difficult to design in from the start…

There may be other issues, but the two biggest ones have to do with the trajectory, and it might be possible to design the trajectory to avoid them.

Enabling Technologies
This approach shares many of the enabling technologies with the other two approaches. Reusable TPS, orbital tugs to offload some of the dead-weight on the stage, suborbital vehicles help provide experience with handling similar vehicles, composite tanks always help with HTHL design (since you can now do a cryogenic “wet-wing”, and have more integrated structural tankage/insulation), etc.

There’s another potential non-technological (regulatory) enabler that an affiliate of ours at MSS is working on, but I’m not sure if I can go into it yet. It would also be beneficial to suborbital operations including both HTHL and VTVL operators.

The Path Forward
As you’ll be noticing if you’ve read the previous parts, there’s a common theme for most of these orbital RLV approaches. Almost all of them have big unknowns when it comes to TPS. Almost all of them can benefit from work being done currently for suborbital vehicles. Most of them can benefit from subscale “proof-of-concept” testing using suborbital vehicles in development as “first stages”. This is particularly the case for this approach.

In fact, the HTHL work that XCOR aerospace is doing right now for their Lynx vehicle is directly applicable to what would be needed for a glideback TSTO design. In fact, as they’ve been saying for a long time, they’re planning on using Lynx or Lynx Mk2 as a nanosat launcher. Using a slightly modified Lynx or Lynx Mk2, you could do work on things like staging techniques, trying out various trajectories, abort mode practice and planning, etc. Not to mention that the technologies being developed for Lynx and Lynx Mk2 (especially the cryogenic LOX tanks) are directly relevant to this TSTO approach, for the exact same reasons. I know that the XCOR guys, for good reason, are very quiet about their ideas about how to proceed beyond suborbital, but I’m almost positive that something like this is how they’d go about it if they were ready to take that next step.

But as with the other approaches, while the path ahead is fairly clear, it’s still involved. XCOR’s been doing excellent rocketry work for almost 10 years now, and they’re barely getting enough traction in the funding world to get their suborbital vehicle into full-time development. But once it’s in operations, taking the next logical steps should be relatively quick–provided someone has the funding and the interest. But that’s a post for another day.

In the previous installation, I discussed approaches to incrementally make ELVs more reusable (or at least recoverable/refurbishable). I discussed why I think that while making ELVs recoverable will be an improvement over the state of the art, such incremental improvements may actually be on a different evolutionary path from high-flight rate capable, truly reusable launch systems. I then discussed the key challenge for TSTO RLVs: how to get the first stage back after a mission, and I outlined the benefit of having the first stage be able to return itself to the original launch site without having to land downrange. This article and the next several in the series will focus on TSTO approaches that provide for return to launch site capabilities.

The first of these approaches, what I like to call “Pop-up TSTO”, has gained quite a bit of attention over the last several years, particularly due to Patrick Stiennon and David Hoerr’s book “The Rocket Company” (which they had me review here, and here). The basic concept is to have a TSTO vehicle, where the first stage flies up purely vertically (John Carmack, who is a fan of the approach has likened the first stage in this concept to a freight elevator) with an apogee of around 100km, the second stage separates from the first stage, and then the second stage provides all of the horizontal acceleration to reach orbital velocity. The first stage reenters and lands vertically like the suborbital vehicles that we at MSS, as well as our friends at Armadillo Aerospace, TGV, and Blue Origin are trying to do. The upper stage after delivering its passengers or payload, reenters and also lands at the launch site.

Benefits of the Pop-Up TSTO Approach
There have been several benefits posited for this TSTO approach:

1. The vehicle is very operationally simple. The first stage goes straight up, the second stage straight over. You have at most four important engine ignition events (liftoff, 2nd stage ignition, 1st stage landing, and upper stage landing if the upper stage uses powered landing).
2. If the upper stage T/W ratio is high enough (approximately 1.4) or if the first stage staging altitude is high enough, the first stage ends up soaking up most or all of the typically 1600m/s of losses that an SSTO design would face. This means that the upper stage only has to provide the ~7800m/s needed for orbital velocity, minus ~325-465m/s for the rotational velocity of the earth depending on launch site latitude, yielding a required delta-V of around 7400m/s for most US launch sites.
3. The upper stage main propulsion system only has to operate in vacuum, so all of the engines can be vacuum optimized, giving much higher mission averaged-Isp.
4. The upper stage also doesn’t operate for the most part inside the atmosphere, so it might not need slosh baffles (or if it does, they probably don’t have to be as heavy as baffles needed on a lower stage). It also probably doesn’t need anywhere near as much gimbal authority as a 1st stage would.
5. Staging can be done at high enough altitude that it is a very low dynamic pressure event. Part of what caused the loss of the last Falcon I flight was that the staging ended up occurring at a lower altitude than planned, which imparted higher aerodynamic forces on the stages, which caused a collision between the 2nd stage nozzle and the first stage.
6. The first stage ends up having performance requirements more like a suborbital launch vehicle than a typical orbital first stage. This means that it’s easier to make it robust and simple, costs can be lowered at times by throwing weight at problems (since the first stage is very weight insensitive). This also means that the first stage could be either evolved from a future suborbital launch vehicle, or at least could possibly be developed by a team that has worked out the challenges of a VTVL suborbital vehicle.
7. Since the upper stage has such a high delta-V requirement, it will end up having a relatively high propellant mass fraction, which means that when it reenters, it will be mostly empty and will thus be very fluffy. Having a low ballistic coefficient (ie a low mass per unit frontal area) means that you decelerate quicker, higher in the atmosphere where the density is lower–this yields both a lower peak g-loading, but also a lower heat flux, thus making the TPS material challenge somewhat easier than for a dense reentry vehicle like the shuttle or most capsules.
8. Since the first stage has no downrange velocity, it’s Instantaneous Impact Point stays right around the launch site throughout the flight. This makes it easier to launch over land, out of more populated areas (instead of having to launch along the coasts or from islands or sea platforms out in the ocean). Most of the high-risk phases of flight (ignition, max-Q, staging, upper stage ignition, etc.) happen when the IIP is within spaceport grounds, and thus away from the uninvolved public. This should make it easier to get licenses for the vehicle to operate out of less traditional launch facilities, which may be a key to lowering some of the cost of space access–and to being able to get more customers for said vehicle.

Now, there are probably other advantages, but those are some of the primary ones as I see it.

Challenges, Constraints, Limitations and Drawbacks
Like with the Air-Launched “Assisted SSTO” concept I discussed in Part I, the Pop-up TSTO approach does not come without its own set of problems. There are always both pluses and minuses to all approaches, and the key to good engineering is to make sure you understand what those limitations really are so they can be dealt with properly. Here are a few of the main drawbacks that stick out to me:

1. Much like the air-launched SSTO rocket stage, the upper stage for a Pop-up TSTO vehicle still faces a nearly-SSTO level of delta-V requirements. Due to the non-linearity of the rocket equation, knocking off 1600m/s vs. a ground launched SSTO makes a huge difference, but providing 7400m/s in a single, reusable stage is still challenging.

   As an aside, many commenters on my air launched SSTO concept seemed to think that such a vehicle wasn’t really technologically doable, but that a Pop-up TSTO stage would be relatively easy to build. I stayed up till 2am doing the math last night, and the fact is that the two are not as different as you might think (I can provide some of the math and explanations if people are interested). The Air-launched SSTO stage needs about 8000m/s (maybe 100-150m/s less for a stage using a more dense propellant combination, or one that has a high thrust to weight at ignition due to using Thrust Augmented Nozzles), compared to 7400m/s for the Pop-up TSTO upper stage. What this equates out to is that for two stages using similar propellant types and similar propellant loads, the pop-up upper stage would only have 20-25% more mass to play with than the air-launched SSTO stage. Specifically for a LOX/LH2 upper stage, you’re talking about propellant mass fractions (the propellant mass divided by the stage plus payload mass) in the range of 0.81-0.82 for the pop-up stage, and around 0.84 for the air launched stage. For LOX/HC, the numbers are around 0.89-0.91 for the pop-up stage, and and 0.9-0.92 for the air launched stage. While that 20-25% more dry mass is nothing to sneeze at, it’s a lot closer than most people would seem to believe.

2. The upper stage needs a relatively high stage thrust to weight ratio at ignition in order to avoid incurring drag losses (around 1.4 being ideal according to The Rocket Company). While you could theoretically loft the first stage a bit higher to give more time, this quickly starts putting your abort g-loads in the range that is problematic for manned flights. So, you either end up taking a small delta-V hit (thus pushing you closer to the air-launched SSTO case), or you end up taking a mass ratio hit for larger engines.
3. The upper stage ends up being very sensitive to weight growth. Adding 1 pound to the upper stage could require an additional 20-30lb worth of hardware and propellants on the first stage. This either means designing in lots of performance margin on the first stage, taking a hit to payload, having to spend a lot more money on weight control on the upper stage, or possibly all of the above.
4. The high delta-V requirements, and the sensitivity of first stage weight to upper stage weight growth push you towards LOX/LH2 or at least LOX and one of the lighter hydrocarbons (cryogenic methane or subcooled propane) for the upper stage. This is typically done by the ELV people as well, but the complexity of adding a cryogenic fuel on-board is annoying.
5. The typical configuration for a pop-up TSTO is going to be two serially stacked stages, which now requires ground handling equipment for stacking stages. This costs money and makes it harder for a given location to setup a launch site.
6. Because the delta-V split on the stages is less than optimal, this results in very big first stages (depending on the achievable propellant mass fractions). Which means that as you scale up, at some point you’ll wind up with a stage that’s too big for normal ground transportation. And because RLVs will typically have a much lower payload to GLOW ratio than ELVs, you’ll run into this road/rail transportability limit at much smaller payloads than ELVs do. For instance, if you don’t go with a LOX/LH2 upper stage, even a very light RLV (1-2klb payload) could end up having a first stage that’s as big as a Falcon IX first stage.

   There is one possible work-around to that problem–and that’s having the first stage be modularly assembleable. While I think John takes the modularity concept way too far (I’d never go more than 7, and would generally try to keep it to 3-4 parallel units), and while I’d definitely go with a more aerodynamic module configuration with higher aspect ratio modules than he has, modularity could possibly help with getting around this problem. Think Saturn-IB first stage except having the separate tanks modularly assembleable, instead of preassembled. Sure, it’ll cost you a lot more integration, and a lot more mass for the mechanical, fluid, and electrical interconnects, but your first stage is already fairly weight insensitive. This would allow you to scale up by at least another half order of magnitude, and by that point you’re probably up into the light EELV range–which RLVs won’t be approaching in the near term anyway.

7. You’ve still got to deal with TPS for the orbital stage.
8. Because the most likely configuration for a pop-up vehicle is two vertically-stacked stages, the upper stage may need to be able to separate itself from the lower stage in some abort modes. While HTHL vehicles can more readily survive propulsion failures at most points in their flight, VTVL vehicles like the pop-up TSTO would likely be don’t have the option of just dumping most propellants and gliding down to an emergency landing. If you have a full propulsion failure of the first stage, it may require separating the upper stage in a hurry. Since this a reusable stage though, typical expendable launch towers aren’t a practical answer, which involves some sort of reusable escape engines (possibly an aggressive TAN extension to the upper stage primary propulsion system). Testing these and making these abort modes safe and graceful is going to be non-trivial.

Enabling Technologies
Being a less aggressive design approach than the Air Launched SSTO, there aren’t as many enabling technologies that aren’t already on the shelf. Thrust augmentation could possibly be helpful (especially for emergency abort operations), but aren’t necessarily required. Composite propellant tanks and structures could reduce the weight of the upper stage a bit, making it easier to hit mass targets, but the upper stage is probably within the realm of feasibility even using metal tanks and existing manufacturing processes. The first stage development and operations would benefit from the existence and flight experience provided by suborbital VTVL RLVs.

The main enabling technology for this style of RLV is going to be the TPS system (and possibly the reentry technique). There are a couple of interesting options out there that might be doable with such a fluffy reentry stage, such as metallic TPS like was planned for Dynosoar or X-33. And there are some more exotic ideas I’ve heard such as Joe Carroll’s “spike” idea. But the reality is that none of these have been proven out yet, and that’s the only real enabling technology for Pop-up TSTOs that isn’t already on the shelf. It’s important to note that this is the case for all of the RLV techniques I’ll be talking about. There are tons of good ideas, but very limited flight data.

Also, looking back at what I said in Part I, all RLVs could benefit from commercially available prox-ops tugs.

Remaining Unknowns and the Path Forward
Unlike Air-launched SSTOs, there are far fewer unknowns that I can see for this approach. The upper stage is still fairly aggressive, so there’s some questions about if we can make a highly reusable stage with the required performance. There’s still the questions about the TPS. And the other big unknown is going to be how to handle aborts throughout the flight regime. In order for an RLV to make economic sense, you can’t be losing it frequently. Just getting the crew, passengers, and/or cargo out isn’t enough if you can help it. Figuring out how to design a reliable VTVL vehicle that can survive reasonable failures is going to be a challenging task. And figuring out how to perform a rapid separation in possibly adverse conditions without adding so much mass or complexity to your upper stage that you make the vehicle less reliable or unworkable is also going to take work.

The key to moving forward though I think is pretty clear. VTVL RLV companies like us at MSS and our friends at Armadillo and the others need to keep plugging along until we are actually reaching 100km on a repeatable and affordable basis. We need to keep working our way up the learning curve, and hopefully finding businesses along the way to make that possible. Once we’re there (or possibly sooner if XCOR or Virgin beats us to space–which is actually fairly likely), subscale TPS experiments need to be done using suborbital vehicles. This can be done using the “nanosat launcher” suborbital RLV upper stage I mentioned in Part I. By decreasing the cost of actually getting real flight data into the hundreds of thousands range might allow for enough iterations to work out some of the bugs on the small scale before trying to build a full-scale prototype. Also, once a VTVL suborbital vehicle is there, most of us in the industry plan on trying to use our vehicles as a first stage for launching nano-sats. This should help work out the challenges of stage integration, staging, and could even provide an environment for testing out subscale launch escape systems and techniques.

Once all of the subscale work has been proven out with suborbital vehicles, it should be much easier to start into developing a prototype orbital vehicle. There’ll still be a lot of work involved, and there will still be some scaling risks, but by using suborbital vehicles to prove out the various concepts, a lot of the important risks can be retired before its time to start work on a full-up orbital RLV.

This article will focus on the key tradeoff that stems from this question: whether to try and recover the first stage downrange, or whether to try and perform some sort of return to launch site maneuver. The answer to this question is probably the number one driver of what approach one takes for developing a TSTO vehicle.

RTLS vs. Downrange Recovery
As I pointed out in my brief discussion about SSTO vs. TSTO approaches in Part I of this series, attaining orbit is mostly about building up a lot of horizontal velocity, and only a little bit about gaining vertical altitude. For performance optimized TSTO ELVs, the first stage often imparts a significant portion of the overall delta-V (especially for ELVs delivering satellites to GTO or GSO). This means that it ends up coming in hot, fast, and a long way downrange from the original launch site. Now, there are several different approaches to deal with this problem (or avoid it altogether).

One option is to just let the stage come down where it wants to, and recover it downrange. Downrange recovery can take several forms including recovering a stage out of the ocean after a splashdown, landing the stage at a downrange site and ferrying it back (either by rocket flight, a carrier plane, or by truck, train, or barge), or it could involve mid-air recovery of part or all of the first stage. While downrange recovery may is the general approach that probably imposes the smallest performance penalty, each of the actual approaches to down-range recovery have some pluses and minuses.

Splashdown Recovery
Let’s take splashdown recovery first. Falcon-1 is an example of the splashdown recovery. The stage separates where a typical ELV would want to have a staging event, and then (hopefully) it’s fished out of the ocean and refurbished for reuse. Some of the benefits of splashdown recovery:

1. Splashdown recovery is probably one of the easiest and best understood methods for recovering a traditional ELV-like first stage.
2. There’s a large experience base to use as a foundation for carrying out such a design.
3. Even if your flight rate is low enough that it isn’t saving you much money, you’re still able to learn a lot from being able to perform post-flight inspection on the propulsion hardware. Thus, even if you aren’t flying enough to save a lot of money via recovery, it will help your reliability.
4. Ocean splashdowns don’t require anywhere near as heavy of recovery equipment as land parachute landings.

But they also have several drawbacks:

1. Trying to make a complicated rocket engine sea-water compatible, especially a turbopump-fed rocket engine, is not a trivial task. Material selection, and getting the stage out of the salt water (and cleaned out) as quick as possible are all required.
2. There’s a lot of time and labor involved in hauling the stage back, cleaning it out, making sure nothing got damaged on reentry or splashdown, testing everything to make sure it’s still in working order, etc. This fundamentally limits how frequently you can refly a given stage. It also translates into a lot of extra personnel and labor-hours required above and beyond what you would normally need to just build, test, and fly an expendable vehicle.
3. The wear and tear from ocean recovery, splash down, etc. are likely going to limit the number of reflights you can get on a stage or engine before major overhaul or outright replacement.
4. Your potential launch sites are limited, since you need a large body of water on which you can drop big heavy hardware. Most likely (for US entities) that means flying out of one of the existing ranges like Wallops, Vandenberg, or Canaveral. These locations, while excellent for flying missiles, and while also improving their commercial friendliness over time, are still a long way from the environment you want to be operating a reusable launch vehicle out of.
5. While it’s possible to design a launch vehicle splashdown recovery first stage in such a way that a first stage failure doesn’t necessarily imply the loss of your cargo, it is much harder to design such systems for graceful abort modes. Unless the upper stage is also designed for splashdown recovery (with the payload designed for it as well), a stage failure probably will result in loss of payload. This loses you one of the big potential advantages of reusability–graceful and intact aborts.

Mid-Air Recovery
The idea behind mid-air recovery is that instead of allowing the stage to crash down into the water, you instead snatch it (or a high-value part of it) out of the air using a helicopter or other sort of aircraft aircraft. This is similar to how Genesis was supposed to be recovered, and was the method used for recovering a lot of the film capsules from early spysats. There are actual serious players looking at this idea, but I don’t know if it’s supposed to be public knowledge yet, so that will have to be a post for another day. There was also a paper floating around by a company that does mid-air recovery work, including work for the SpaceHab ARCTUS project. If I can dig it up again, I’ll probably post about that as well.

Anyhow, here are some of the benefits of mid-air recovery:

1. No salt water contamination in the rocket hardware! This greatly cuts down on the amount of work that needs to be done to turn a stage around. No need for decontamination. No need for stripping down hardware. Probably eliminates the need to “requalify” the propulsion system before reflight.
2. Gentle, low-shock recovery is much less likely to damage stage or propulsion hardware, also making it more likely that the hardware can just be reused after some inspection.
3. There are companies that specialize in this sort of thing, and you can just rent their services instead of trying to do this in-house. They aren’t cheap, but they’re a lot cheaper than building a new stage every time.
4. Your propulsion system is going to be in about as close to the same condition as it was when the engine shut down as you’ll get for any recovery technique–this makes it a lot easier to get good reliable data on wear-and-tear on the engine, so you can improve the quality over time.

But here are some of the challenges:

1. Complex recovery technique. Sure, you can practice it a whole lot for not too much money, but there is some increased risk of failing with the rendezvous or recovery operations, which could occasionally cost you a stage.
2. Weight limits. Even with the latest techniques, which can recover payloads up to 80% of the maximum cargo capacity of the helicopters, you’re still limited to around 22klb or less. Depending on the size of your stage, this may mean that you can only recover part of the stage (like say the engines). That’ll still likely save at least some money, but it’s not as big of a win as getting the whole stage back intact.
3. There may also be issues with trying to recover a big, but fluffy stage. Depending on the weight distribution, there could be some real oscillation issues (like what happened when they tried to move the Roton ATV under helicopter).
4. Range issues. Depending on how far downrange your stage comes back, you might need to also rent not just a helicopter, but some sort of barge to operate the helicopter off of. This will increase the amount of time it takes to turn a stage compared to if you could just fly it back.
5. Like with splashdown recovery, this method of recovery still doesn’t give you graceful and intact recovery methods in the case of a first-stage failure. With dump valves and two helicopters, and a mid-air recoverable upper stage, you might be able to recover the payload over part of the trajectory, but you’ll still have zones where a failure means sure loss of the payload (or a launch escape abort if you’re flying people). It isn’t a showstopper, but it does reduce the upside somewhat.
6. Due to challenges #1 and #5, you probably still need to launch out over the ocean, which means that once again you’re still going to face the issue of launching out of an existing missile range. Basically, since there’s a chance you could biff the in-air recovery, you have to do this over an unpopulated area. And since your vehicle doesn’t likely have graceful failure modes, it’s more like an existing ELV than a more traditional RLV, and will probably be treated as such by the FAA and the ranges. Not a showstopper either, and it might just be possible to pull this off with an over-land launch if you can find a sufficiently deserted area, but definitely a challenge.

Mid-air recovery is probably too weight constrained for something like a complete (but dry) Falcon IX first stage, but might be an interesting option for recovering the Falcon IX upper stage or the Falcon I first stage. It’d also probably be just the right size for recovering the first stage if they hadn’t canceled the Falcon V. Other than the weight limit, there’s some real benefits of this approach over the traditional splashdown technique.

Downrange On-Land Recovery
This type of recovery can take several forms. It could be a powered VTVL landing at a downrange pad. It could be a powered or glide landing for a HTHL. It could be a parachute and airbags landing (like Kistler, just downrange). But basically you have the thing land, on the land, downrange, and then fly the thing back, or ship it back.

Here are some of the benefits:

1. Much more efficient, performance-wise, than any of the RTLS approaches. You can still stage at the most optimal staging velocity, therefore making your upper stage design a lot easier. You also get a lot more payload per given takeoff (and dry) mass.
2. At least some of the RTLS approaches can also sometimes use this as a performance enhancing option–in case you need to launch a bigger payload than you can handle with a normal RTLS trajectory.
3. Unlike mid-air recovery, this recovery approach can scale up to fairly large sizes.
4. In emergency cases for RTLS approaches, you may want to be able to land your vehicle at alternative downrange sites anyway.
5. Unlike the other two downrange recovery options, this option is a lot more compatible with intact and graceful aborts.

And here are some of the challenges:

1. A given launch site will typically have its launch azimuths (directions in which you can launch) restricted a lot more for downrange land recovery than it will for an RTLS vehicle. This is because you need to have a suitable place downrange where you can actually land. This makes downrange recovery vehicles less flexible than RTLS capable vehicles.
2. You need facilities at both ends, especially if you intend to fly the stage back after landing.
   This may entail having almost as many launch support people at the downrange site as at the initial site, which greatly increases the fixed costs of such a system. Probably not quite double (since you don’t have payload processing facilities there), but it’s a non-trivial expense.
3. If you do a rocket powered return, you’ve now effectively halved both your flight rate (as you have to do two launches, two landings, two ground preps, etc. per a single paying flight), and halved the number of revenue generating flights you can get out of a given airframe. Both of these directly affect the bottom line.
4. If the return flight is a rocket-powered suborbital flight (as per AST’s definitions), I think that each of your downrange sites will need to be an FAA licensed launch site, and you will need launch licenses for all of the return flights. Now, once you have one launch license to base things off of, getting additional ones should be easier, but its still extra paperwork. Also, your Ec and MPL calculations are going to be different for the return flight, because your IIP will move at different rates over different areas under your groundtrack for the two trajectories (not to mention mission-critical operations will occur with your IIP over a different location). All of this stuff has to be taken into account.
5. If you have a jet powered return (either using a carrier aircraft, or if the stage has built-in jet engines), you now need to deal with the aircraft side of FAA, which may entail getting the vehicle type-certified. I’m not certain, but having a vehicle that operates under both regimes is likely going to make things a lot harder, not easier. Being unusual is not a virtue when dealing with regulators. If you’re using an existing carrier craft, that’ll make things easier however, as it is purely a subsonic aircraft, and thus a lot closer to what FAA is used to dealing with.
6. If you try to return the stage via trucking or train, now the stage has to be “roadable”. Which means making it skinny enough to fit on existing transports. While this may be feasible for some smaller, dense-propellant RLV stages (after all I think that Falcon IX is roadable), it is a constraint on the size of stage. And the aspect ratio roadability forces you into is not as ideal for VTVL stages. VTVL stages want to be shorter and squater than typical rocket stages.
7. If you return the stage via trucking or train, you now need heavy moving equipment at any downrange sites, experienced heavy equipment personnel there, and it’s going to cost you a lot of extra time. All of these things add cost, and slow down your turn time.

Conclusions: The Case for RTLS
Now, I probably ought to clarify something. I don’t think any of these downrange recovery ideas are stupid. If done right, they can save a lot of money compared to a purely expendable system, while also increasing reliability by allowing for post-flight inspections and the like. Especially with the downrange land-landing techniques, you can get all of the benefits of traditional RLVs.

In other words, while there are some challenges with downrange recovery, there are often some real benefits. There are some cases where using these downrange recovery approaches really is the best choice. SpaceX and the others looking at these approaches aren’t being foolish by pursuing them. I just think that the inherent limitations of this sort of reuse (especially the first two options–splashdown and mid-air recovery) will probably prevent it from being a revolutionary as opposed to a modest, evolutionary improvement over a purely ELV approach. Now, in the near to medium term, even when RLVs first start flying, they’ll likely be relatively quite small compared to the EELVs (for reasons I’ll go into in a later post). Which means that approaches that allow existing ELVs to become somewhat more reusable, and improve their economics somewhat are actually useful. I think that while small RLVs will bump ELVs out of the people, light cargo, and propellant markets very quickly after they enter the field, there’ll still be payloads that are too big for RLVs that are small enough to be economically viable in the near-to-medium term. The ULA’s, SpaceX’s, and Sea Launch’s of the world will still have a useful role to play for some time yet. Particularly for launching bigger payloads like Bigelow stations, transfer stages, etc. So, having ways to improve them is good, even if they aren’t necessarily going to change the world all by themselves.

As for the last options–land recovery downrange, it actually does have the potential to be revolutionary. But the approach still has some serious economic and regulatory drawbacks that are sufficient to make one start looking at RTLS approaches, even though they may be less optimal from a purely performance-based standpoint. There are four primary (and one somewhat oddball) RTLS techniques/trajectories: pop-up, glideback, boostback, flyback, and once around return. Of these five approaches, the most well known (until recently) and thus most thoroughly studied is the flyback approach. However, the first three are the ones (pop-up, glideback, and boostback) that I think are the most promising and relevant to near-term orbital RLV endeavors, and thus will get the bulk of my focus for the remainder of this series (Parts III-V). But I’ll probably spill a few electrons discussing the last two as well. They are interesting after all.

The first such approach I would like to introduce for discussion is epitomized by a proposed design (illustrated below, credit: Teledyne Brown) that was brought to my attention about a year ago. This proposed design, termed “Spaceplane” was developed at Teledyne Brown by Dan DeLong (who later became one of the founders XCOR Aerospace and is currently their Vice President and Chief Engineer, and who also currently owns all the rights to the Spaceplane design). Dan’s proposed concept was a winged, “assisted” single-stage to orbit (SSTO) design that was launched off of the back of a converted 747. The LOX/LH2 stage, powered by 1x SSME and 6xRL-10s would theoretically be capable of delivering ~14klb of unmanned cargo to a 400km circular orbit. The vehicle would be reusable, using an Inconel-foil over fiberglass insulation concept for its reentry TPS, and using a runway landing for its recovery method.

While the specifics of Dan’s proposed design are now a bit dated (the concept was proposed back in the late 80′s), the general approach still merits investigation.

To Stage or Not to Stage: That Is The Question

Now, before I go into the specifics of this approach, I know at least a few of you are probably already thinking things along the line of “SSTO? He can’t be serious. Everyone knows that SSTOs are totally unrealistic!” While to be honest, I’m mostly a TSTO guy myself (as is Dan DeLong these days), but I think there’s a real danger in how quickly and without contemplation people tend to buy into new conventional wisdoms.

The fundamental reason why anyone would even want to stage a rocket vehicle has to do with the physics of the rocket-powered flight. The rocket equation, says that the change in velocity due to a rocket in flight is linearly proportional to the specific impulse of the propulsion system and proportional to the natural logarithm of the vehicle’s mass ratio (the ratio of the mass at ignition to the mass at shutdown of the engines).

Last week, I was asked to do a remote guest lecture for a university course on space development (being run by Dr Livingston). It was somewhat flattering to be grouped in the same category as much more experienced space technologists, pundits, and businessmen such as Dennis Wingo, Michael Kelly, Jeff Foust, and others. As part of the presentation on developing reusable orbital transportation, I discussed a short list of orbital space transportation approaches that I felt were the most promising directions for development.

So, over the next several weeks, I want to take a little bit of time to introduce and discuss some of those proposed approaches for reusable orbital transportation. Now, a lot of this may be a boring rehash for fellow engineers and technologists, but hopefully I can provide some useful discussion for those coming to this industry from non-engineering backgrounds. I’m planning on discussing the basic concept behind each approach, the potential pros and cons, the unknowns that need resolving for said approaches, and some thoughts on incremental development methods for resolving those unknowns. I may also go into some of the other topics I discussed such as my ideas on reusable transportation markets.

My goal is to provide a basic understanding of where we are, what we think some potential solutions might look like, and an understanding of some of the more probable paths that could take us from here to there (technologically). With that information as a background, it will hopefully make it easier to discuss how business, financing, and government policy issues tie in with the technological situation.

Hopefully I’m not biting off more than I can chew.