Some Interesting Ideas From the Other Side of the Pond

I don't have time to go into detail at the moment, but I wanted to relay an interesting paper that Keith Cowing reported on NASAWatch today. Now, if I were someone at the ESA, I'd probably be taking NASA's grand plans about Constellation with an appropriate sized grain of salt right about now. But there were some good ideas overall:

• The report mentioned that our ISS experience shows the importance of having redundant transportation methods (ie imagine what would've happened to ISS if Soyuz didn't exist). I don't think that redundant transportation method should necessarily be another government-centric transportation system, but I agree wholeheartedly that monocultures are a bad idea.
• The report also mentioned that having a safe-haven in LLO is one of the best ways to increase the safety and flexibility of a lunar exploration program. Right now, most of the danger associated with lunar exploration have to do with operations on or near the moon. The current architecture does nothing to reduce those risks, but instead focuses on the much sexier earth-to-orbit transportation risks. Having some infrastructure in LLO can go a long way to fixing that, while also giving you some very interesting mission options. Now, I'm still a fan of the idea of Lagrange stations, and I think that in the long-run they'll dominate the traffic in the lunar half of cislunar space. I just think that there is a small, and critical niche filled by one or more small polar LLO stations. I've been planning to write up my ideas on this concept for over two months now, so can someone poke me in a few weeks if I haven't followed up on this thought?
• Unlike NASA they don't seem to be deathly afraid of on-orbit assembly when it makes sense. Of course, they don't have an HLV fetish that they have to rationalize...
  

There were a few other good points, but those three were the key ones that stood out to me. Of course they also seem to be missing the importance of propellant transfer, and they seem to be almost as clueless as NASA as far as commercial enterprise is concerned (both why it's important, and how best to foster real commercial involvment). But it was an interesting read if you have a few minutes.

Orbital Access Methodologies Part V: Boostback TSTO

While I have the topic fresh in my mind, I decided to jump into the next part of my continuing series. Though it wasn't a conscious choice on my part, I notice that the order I went with for this series actually follows a consistent pattern. In each part of this series, we discuss methods that move more and more of the delta-V load off of the orbital stage and onto the carrier vehicle or the first stage. In the case of Air-Launched SSTO, the carrier plane removed about 1000m/s from the ~9km/s normally required for a ground-launched SSTO, thus making an SSTO design feasible. For the Pop-up TSTO design, the first stage's vertical trajectory removes all of the gravity and drag losses from the upper stage (a savings of ~1.6km/s). For the Glideback TSTO design, by using aerodynamic lift to turn around and glide back to the launch site, some horizontal downrange velocity was added, thus lowering the delta-V requirements even further (probably saving somewhere between 1.6-2.4km/s depending on the details). The next approach we'll discuss follows this same trend.

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.

Orbital Access Methodologies Part IV: Glideback TSTO

Some of the comments to my last post got me thinking about what I'm trying to accomplish with this series. The reality is that each of these approaches that I'm discussing could easily fill a full chapter in a textbook, complete with 20-30 pages of text, tons of graphs, equations, sample designs, detailed discussions of tradeoffs, etc. I'm probably not the guy you would want writing such a textbook--that's something better left to either a Masters/PhD student looking for a fun dissertation, or someone who has more aerospace engineering experience than myself (say a Mike Kelley, or a Dan DeLong or maybe a group of such people).

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.

Orbital Access Methodologies Part III: Pop-up TSTO

This third installation in my Orbital Access Methodologies series (parts I can be found here, and part II here) has been a long time in the coming. It has taken so long, not because I've been spending months researching and analyzing the topic (I knew most of what I wanted to say back in January), but mostly because I was surprised by how much favorable attention the first part received, and I've been worried about not meeting expectations. A good part of the reason why that first article was so good was that I was able to lean heavily on help provided by Dan DeLong and Antonio Elias, both of who had been analyzing air-launched orbital access methodologies since I was still in gradeschool. I now have a bit more empathy for movie directors trying to make a sequel or a prequel to a first movie that had been far more successful than they had ever thought.

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.

Westward Ho?

Other than a busy schedule at work over the past several weeks, and ongoing blogger's cramp, one of the other big reasons why I haven't been blogging very much lately is that Tiff and I started reading together again. This is an old tradition of ours that we started back when we were poor newlyweds and couldn't find a cheaper date than borrowing a book for free from the library, snuggling up, and reading to each other. Anyhow, when we finished the Harry Potter series a few months back, I had decided I needed a break from reading together every night, so I could get caught up on my blogging. I still have another one or two installments to write in my Orbital Access Methodologies series. Seeing as how that hasn't really been happening, we decided to finish up the last two books in a nine-book historical fiction series we had started back in Santa Clara. To my surprise, reading this series has actually got me thinking more about space development, and so I figured I'd share some of my thoughts.

The series, The Work and The Glory, is a historical fiction adaptation of LDS Church history over the 1828-1847 timeframe, revolving around the lives of a fictional family, the Steeds. While the books do tend to get somewhat preachy at times, and while someone familiar with LDS history might find some of the foreshadowing to be a little weak, the series was overall a good read.

The Mormon Exodus
It was the last two books in the series that got me thinking about the challenges of space settlement. These two books, which we finished a few days ago, cover the Mormon exodus to Utah in 1846-47. In 1845, the deteriorating relationship between the Saints in Nauvoo, Illinois and their non-LDS neighbors got bad enough that the Saints agreed to leave the state by the end of the following Spring. Since none of the other states in the Union at the time were willing to accept the Mormon refugees, Brigham Young and other church leaders decided to colonize the Great Basin in the heart of the Rocky Mountains (which at the time were part of Mexican territory). The theory being that the area was pretty much unpopulated, nowhere near as nice as Oregon and California, and therefore they might finally be left alone.

What the story really brought out was how amazingly challenging of an undertaking the exodus was. The goal was to move all 15,000 Saints (many of them destitute) over 1200 miles through unsettled wilderness, and set up civilization in the previously unpopulated mountain valley around the Great Salt Lake. Much like space, the destination was not existing towns or settlements, and relative to other areas being colonized at the time, the area was very forbidding and unfriendly to human habitation.

One of the interesting points I gleaned from the story was the importance of infrastructure in making it possible to move that many people, and Brigham's use of advanced teams to help prepare the way for the main body of settlers. The original plan in late 1845 and early 1846, was that a set of advanced companies would lead-out, with the goal of reaching the Valley early enough that summer/fall to plant some quick-growing crops. The hope was that they could get there early enough to provide food for the main body of the Saints to survive the winter by the time they would arrive. Also along the way, they'd be blazing the trail: creating fords, digging down steep river banks to allow for easier crossing, building bridges or ferries where necessary. As the nasty weather that Spring in Iowa Territory bogged the Saints down, these advanced parties were instructed to build temporary settlements at several locations including Mount Pisgah, Garden Grove, and eventually Council Bluffs and Winter Quarters (near modern-day Omaha, Nebraska). They fenced in and cleared farm land, planted and cultivated crops, and then moved on, leaving those crops to be harvested by those who were still coming up the trail.

One of the other major lessons I learned was that settlement by large groups is inherently more complicated than settlement by smaller, individual groups. Especially when you need to move not only the rich and well-equipped, but also the penniless, starving, and destitute. In the end, by the end of 1847, only about 10-15% of the Saints had arrived in Utah. Things ended up taking well over twice as long in the end, but created an infrastructure that allowed everyone, no matter how poor to make the trip. Of the ~60,000 people who made the trip before the railroads reached Utah, many of those literally pulled pulled their way to Utah using handcarts. Without the ferries, resupply settlements, trails, rescue parties, the Perpetual Emigration Fund, and other pieces of infrastructure set in place during the initial exodus, many of those who followed would have never made it. Also, the experience gained in setting up the infrastructure for the exodus provided experience that eventually led to over 500 settlements throughout the Rocky Mountains (spanning from Mexico to Canada), playing a pivotal role in the settlement of the Western United States.

Anyhow, there are far more interesting stories (the Mormon Battalion, the Brooklyn Saints, the Donner-Reed Party, etc) that were covered in those two books than I can do justice too. So, if you can stomach the thinly-disguised religious apologetics long enough, I'd highly recommending digging into some of these books.

Important Differences (ie history never really repeats itself even if it rhymes)

While there are many possible similarities between the Mormon settlement of Utah, and the challenge of space settlement in our century, there are also lots of important differences. One of the keys to successfully learning from analogies is recognizing where they break, and understanding how those differences impact your situation. Because, as one of my history professors at BYU pointed out "all analogies fail at some point." Failing to recognize the differences between your analogy and real life leads to the sort of silly debates all too common in the space community.

Some of these differences make space settlement more challenging, while others might make it somewhat easier than the colonization of Utah that we've been discussing.

• At the time of the exodus, the basic technology for traveling directly to the Great Basin existed. Conestoga wagons, draft animals, firearms, ferries, sailing ships, bridges, etc were all technology that was already in use for commercial and military logistics purposes, as well as playing a big role in the settlement in the MidWest and other places. Much of the technology had been around for centuries or millennia. Not only were the technologies well-understood, but they were commercial available, cost-effective, and off-the-shelf. While there were still improvements being made continuously, as far as settlement was concerned, the commercial state-of-practice for ground transportation at the time was far beyond what the commercial state-of-practice is for space transportation today. A particular case in point is the wagons.
  
• Along the trail and at their destination, the Mormon colonists were frequently surrounded by ready sources of food, “fuel”, water, and raw materials for repairs. While it was still possible to starve and die or to run out of water, or to break down in an area where there wasn't readily available wood (just ask the Donners and Reeds), the general availability of easily extracted in-situ resources was much better for the pioneers than it will be for future space settlers. After all, even the air isn't free in space. Sure, there are potentially interesting ISRU technologies out there, but they're nowhere near as mature as cutting wood, shooting buffaloes, carrying water in barrels or canteens, or simply allowing the cattle to graze along the way. Not to mention the fact that you didn't have to carry all the food for your animals for the whole first half of the Mormon trail...
• While they weren't as common as would've been nice, there were several human settlements in existence along the way by that point. Places like Fort Laramie, Fort Bridger, and Independence, Missouri. While goods were expensive, they did allow for restocking at some price of hard-to-replace items.
  
• Transportation physics were completely different. While I don't think that the rocket equation makes inexpensive space travel impossible, it really complicates things, and sure doesn't make it easy. Combine that with the lack of stopping places between here and LEO, and the transportation physics are much less favorable for space settlement.
• On the side of things that are easier with space settlement, the Moon is a much shorter trip than the trip from Nauvoo to Salt Lake City. The fact that you're talking about less than a week worth of travel instead of several months makes a huge difference in some of the required provisions and supplies.
• Modern water and air recycling technologies, when combined with freeze-drying can allow for a much smaller amount of food mass per person per given amount of time--potentially over an order of magnitude less.

The Big picture
Basically when you look at where we are today relative to space settlement, we’re nowhere even close to settling the solar system (not even the moon) as the Saints were to colonizing Utah when they were camped at Sugar Creek, across the river from Nauvoo in the bitter winter of 1846. Our civilization as a whole has never even flown 100 people to orbit in a year, let alone 1000, 3000, or 5000.

Just to give a sense of the scale we're talking about, here are some rough numbers to think about. Suppose it takes at least 6000-7500lb per person (which is probably a very optimistic bare minimum) to settle and survive off-planet. If those numbers are accurate, then in order to settle 15,000 people in space, even just in LEO, you’re talking somewhere around 45,000 tons of material needing to be lifted. Doing that over a 3-5 year period like the first wave of the Mormon Exodus would require 9-15 thousand tons to be lifted to orbit every year. That’s over 100 Ares-V equivalents per year, and several orders of magnitude higher than what has ever been done before. And that’s assuming that they all stop in LEO! Going to the moon would require something like 6-10 times that mass in LEO in order to do that, and Mars or Venus would likely require even more. That gets you to 90-150 thousand tons per year!

At least right now, if some group of 15,000 people were given a similar ultimatum to what the Saints got in 1845, they'd probably be screwed.

ISRU, Infrastructure, and Access
Now, in this article, I'm not going to go into the rationales that could potentially justify settlement on that scale, or even if it is desirable at this point in time, or how soon it would be desirable. I'm just trying to paint a picture of the kind of things that would need to happen in order to make such an exodus feasible in the first place. Think of this section as a sort of roadmap for stuff that would need to happen between now and when space settlement becomes an even remotely realistic possibility.

With that in minde, once you start to realize how mind-bogglingly large the numbers are when you start talking about serious space settlement, you realize that if such a state even is reachable, it will require attacking the problem from multiple directions. There are three different primary areas of development necessary for enabling space settlement: space access, in-situ resource utilization, and in-space transportation technologies and infrastructure.

Even though lower cost, reliable, and very frequent space access is probably the most important step in the near-term (and also the one that has the clearest near-term potential for ROI and thus commercial feasibility), I want to start by talking about ISRU. ISRU helps enable space settlement by reducing the amount of mass you need to ship to orbit in order for someone to settle somewhere in space. Using the historical analogy I started in this post, colonizing Utah would've been far more difficult if all the food, draft animal feed, and construction materials for Salt Lake City had to be hauled from Illinois. In fact, if that were the situation, colonization would've been impossible. ISRU basically allows you to hack the space access (ie earth-to-orbit) transportation problem back down to a size that might be workable.

ISRU covers a wide variety of areas including in-situ propellant production, production of life support gasses and liquids, production of construction materials, farming, and eventually extracting and processing industrial raw materials, and manufacturing finished goods. Many people in many venues have talked about this subject (particularly Peter Kokh in his "Moon Miners Manifesto" newsletter), but I want to put my own spin on it. For enabling settlement, there are some ISRU areas that are higher leverage than others (ie where you get more of a payoff for a smaller initial investment). As I see it, the two highest leverage areas for enabling space settlement are in-situ propellant production and in-situ construction and construction materials extraction/processing. Of the mass you need to take to LEO in order to settle on the Moon, Mars, the upper Venutian atmosphere, or even the asteroids, most of it is going to be propellant for the trip. Eventually, as more advanced in-space transportation technologies get fielded, this may change, but if space settlement occurs in the near to medium future, I think the importance of oxygen and hydrogen (and possibly some light hydrocarbons like methane or propane) is major. The next biggest mass is going to be the actual structures that people live in and the other buildings, roads, etc. In fact, some previous studies have pointed out that if you can use ISRU to provide for spacious extra-terrestrial facilities, it can have spillover effects on lots of other things. If your internal space is relatively spacious, for instance, you might be able to have more of the work you need to do be done inside, in a shirtsleeve environment, thus allowing you to more directly leverage existing terrestrial tools and processes, without having to do as much redesign.

When you look at in-situ propellant production, something you realize very quickly is that the less you have to ship propellants around, the better. In other words, the closer you can harvest propellants to the place they will be used, the better. That's one of the reasons why even though it's a long-shot, I'm so interested in the whole concept of atmospheric propellant gathering. LEO may be halfway to anywhere, but it's also halfway from anywhere too. If you can gather LOX in-situ in LEO, and especially if you can do it in large quantities at reasonable cost, that would have a major impact on the cost of beyond-LEO transportation. On the other hand, another thing you realize when you study the problem is that due to the rocket equation, propellant resupply on the final legs of the trip have a disproportionately large impact on the overall propellant requirements. Being able to ship a Lunar, Martian, or Venutian lander "dry" saves a lot more weight to LEO than just that landing propellant. It also save the propellant needed to ship that propellant to the destination in the first place. So, at least to me, the two highest payoff places for propellant ISRU are in LEO (if possible) and in orbit around the final destination (for planetary destinations).

Other forms of ISRU such as farming, large-scale manufacturing, extraction and processing of industrial metals, etc. all have an impact on the situation, but for the most part they are much lower leverage--as far as space settlement itself is concerned. We might still see some of the metals extraction and processing sooner rather than later if for instance, it turns out that Dennis Wingo's theories about lunar PGMs turns out to be true. But as far as actually getting there and setting up shop, propellant extraction, and extraction of the simplest construction materials is more valuable.

In order to use those ISRU derived propellants, you need more matured in-space transportation technologies and infrastructure. You need propellant depots, you need reusable in-space transportation (as well as reusable landers). You need technologies that make reuse easier such as better aerobraking (which may involve both infrastructure like satellites, and space technology like better reusable TPS, ballutes, etc). You eventually need infrastructure to service and maintain those transportation systems. You'll probably want lots of prox-ops tugs, and you'll eventually want rescue services (possibly provided by some of the prox-ops tugs). Some of this infrastructure wants to be in LEO (in whatever inclinations have enough demand and/or cheap supply to make sense--and probably eventually multiple smaller depots in the same inclination), and some of it in the vicinity of the destination (for the moon this could mean elements in L1, L2, and/or low lunar orbit, for Mars this probably implies stuff in Mars orbit, or possibly on Phobos or Deimos themselves, for Venus you'd be talking about a Venutian orbit). This infrastructure will probably grow "organically" as market and governmental demand for those services grow. It's hard to know in advance what exact mix of propellants, inclinations, number and size of depots, etc will make the most sense--so it will need to be market driven (and yes, government customers are a market too--just a potentially very dysfunctional one that needs to be treated with a lot of care).

The core reality though is that in order to be able to even get to the point where large infrastructure or ISRU development can really take off, the space access (ie earth-to-orbit) transportation situation needs to improve. Even if you're getting all of your TLI and landing propellants from lunar and upper atmospheric sources, you still need to be able to ship vast amounts of material up from earth, and in order for settlement to be feasible it has to be both significantly cheaper than current launch methods allow, but also the sheer quantity of material that needs to be shipped (even with the rosiest of ISRU scenarios) requires a fundamentally different approach to space transportation. While you may be able to do some of the early infrastructure development with existing launch vehicles (tugs and early "pilot-plant" scale propellant depots come to mind), largescale infrastructure and most ISRU other than atmospheric propellant gathering really need lower cost and more frequent transportation.

As Henry Spencer has put on multiple occasions, developing and debugging ISRU on the moon is going to be an involved process, even if it may be a very worthwhile one. The idea that we're going to design a working ISRU plant, ship it out to the moon, set it up with a few robots, and then start pumping out LOX right away is ludicrous. We know some things about the moon, and there are some good ideas on how to solve some of the more pressing problems, but the reality is that the lunar environment cannot be simulated 100% here on earth, and there are going to be plenty of snags, complications, and unexpected events. Developing hardware, materials, design processes, chemical processes, etc that can cope with the local environment is going to require trial and error and probably several people "on the ground". It's going to take time and lots of work to develop spacesuit materials that can handle the dust, making seals and airlocks that do what we want them to do isn't likely going to be one of those things we get right the first time. And especially once you get into things like metals extraction and processing, developing construction techniques, etc. you see more of the same challenges. And the reality is that the same thing probably applies for Mars, or Venus, or the Asteroids, or even to living in orbit or in deep space. In theory there's no difference between theory and practice, but in practice there always is.

Improvements in space access need to not only include cost, but frequency, reliability and sheer volume. When you look at the air transportation industry, you're probably talking about over 10,000 jets flying every day (some of them multiple times per day) from hundreds or maybe even thousands of airfields. With rocketry today we have something like two dozen flights per year from about one dozen or less active sites. If we're going to get to the point where we're shipping hundreds or thousands of people and their goods to orbit there are things that fundamentally need to change. One of the biggest changes isn't just going to reusability, but going to reusable vehicles that can fly from many locations. While what SpaceX is trying to do is technical reusable (or at least recoverable), they're never going to be able to operate out of more than three or four launch sites (Kwaj, Vandenburg, Canaveral, and maybe Wallops). Same applies for some of the reusability ideas I've seen bandied about by other ELV groups. Sure, for larger goods, those types of reusability are an incremental step in the right direction. But in order to get from where we are now to a transportation system that can fly 1000s of people and their goods to orbit every year, ELVs or recoverable ELVs really stop making sense at some point.

In order to get to the point where we could fly that many people and their stuff to orbit every year, not only do you need "reusable" launch vehicles, but they also need to be capable of high flight rates, capable of operating out of many launch sites (including combination airport/spaceports like Mojave), and safe and reliable enough that they can launch at least some of the time over land. In Part I of my Orbital Access Methodologies series, I discussed one such potential approach, in the next part(s) I'll be discussing a few more.

Summary
Getting to a point as a civilization that we're truly ready to start spreading out throughout the solar system is going to be a difficult process. We're nowhere even close to where we need to be, and most of the options being investigated by national governments are pretty much orthogonal to where we need to go if we want to see our civilization become a truly spacefaring one. Getting to there from here will require work on radically improving earth-to-orbit space access, developing in-space transportation technologies and infrastructure, and learning how to tap the resources of the upper atmosphere, the Moon, and other planetary and asteroidal bodies. There is useful work that can be done now on all three of these areas, but we've got a long way to go, and the engineering challenges are very interconnected. Probably the biggest challenge of all is going to be finding a way to craft solid and profitable business cases along the way to fielding these technologies.

Westward Ho? We'll see, but we probably have an even rockier road between us and our destination than the Mormon pioneers did in the spring of 1846.

Are Those Nozzles "Thrust Augmented"?

In this article at AvLeak, there's discussion of an Aerojet LOX/Kero booster engine project called HC Boost (emphasis mine) :

Dubbed HC Boost, the technology development program is aimed at providing an improved, home-grown alternative to the Russian RD-180, the only other viable current-production hydrocarbon rocket engine. Unlike the RD-180, however, the US engine would be designed to be re-usable for up to 100 missions, have up to 15% better performance and would operate for up to 50 missions between engine overhauls.

Now, the RD-180 is a very high performance engine. It's combustion chamber pressures are actually quite a bit higher than the SSME. When you combine this with the statements in this paragraph (my emphasis again):

The last US-designed and produced hydrocarbon engine was the Rocketdyne RS-27, based on 1960s technology and now out of production. The HC Boost engine, on the other hand, is expected to have higher operability, faster turn time, a longer-life thrust chamber, turbopumps and a new design nozzle.

It really does seem to suggest that they might be talking about thrust augmented nozzles. By going with nozzle thrust augmentation, you could get that same high lift-off thrust, and good sea-level Isp without requiring anywhere near as high of a chamber pressure. Which would necessarily lead to less demand on the thrust chamber and pumps.

A similar engine concept I've talked about with a few friends in the propulsion business is if you took the Merlin-1C upgrade that's been hinted at by SpaceX, and applied that extra pump power they're looking at creating to providing higher flow, lower pressure propellants to a TAN section (say by diverting some of that extra high pressure flow into some sort of a jet pump), you'd be able to get away with a much larger expansion ratio nozzle (with good sea-level Isp), while also nearly doubling the thrust. I really like the idea of a Falcon 1f that can put say 5000lb of payload into LEO for little more than the cost of a Falcon 1e....It's a powerful concept, though I would understand if for now a company like SpaceX is focusing on getting Falcon-1 and -9 flying reliably first. First get it working and working reliably, then add the afterburners...

Which reminds me, I need to go back and do a more detailed article on some of the additional concepts and technical information I've found about Thrust Augmented Nozzles at some point.

Space Tugs vs. Space Ferries: A Useful Distinction?

Something that's been bugging me for some time is the confusion surrounding the term "space tug". The term's been used to describe at least two very different ideas for many years now. At NGEC-2, I tried to inject a little clarity into my working group's discussions by drawing the distinction between "tugs" and what I called "ferries", and I was wondering if others thought it was a useful distinction (and if anyone had a less snicker-drawing nickname then "ferries"--you would think the conference took place somewhere near San Francisco or something from all the chuckles that term drew...)

Under my proposed classification scheme, a "space tug" would be a spacecraft of some sort that primarily is used for maneuvering target spacecraft/objects in the near vicinity of a space station or another spacecraft. For instance, the CSI and CSI/SSL systems proposed for COTS 1 and COTS 1.5 would both fall under this category (and Orbital Express would also likely fit under this category). A "space ferry" on the other hand is a spacecraft that hauls other spacecraft, cargo, or people from one orbit to another in a reusable fashion. For instance, CSI's or Space Adventures' respective "Soyuz-Around-the-Moon" concepts would somewhat be examples of a one-use ferry.

Basically tug == prox ops, ferry == large orbit transfers.

Both are very important capabilities, but while they have some overlap in requirements, many of their requirements lead to very divergent capabilities.

Tugs for instance are explicitly designed for proximity operations in mind. A good tug system implementation would likely have one or more robotic arms for better handling, grappling with, and berthing target spacecraft. A tug likely doesn't have a huge amount of propellant on board. Enough to move things around between various low earth orbits, and to maneuver around the station, but total delta-V capability is probably in the low-hundreds of m/s range. Tugs want to be very robust. The very low delta-V requirements actually make a tug very mass insensitive. So long as most of the things you're moving around are an order of magnitude or more bigger than you, even doubling the mass of your tug has only a minor effect on the total propellant used for tug operations.

Ferries on the other hand are high-performance spacecraft. The delta-Vs necessary for a useful space ferry are on the order of 4-8km/s (though those last 4km/s are probably going to be "dead heading" ie. flying the ferry back to LEO with no payload attached). In the case of a chemically fueled ferry, this means it looks very similar to an upper stage--mostly take, one or two big engines, and some hardware on both ends. An inflatable aerobrake might not be a bad idea depending on how much it weighs. It might not really need much in the way of prox ops capabilities, just navigation and rendezvous capabilities. Ferries are typically going to be much bigger than their cargoes, while tugs will typically be much smaller.

Both ideas also provide different benefits.

The key benefit of tugs is that they enable launch vehicles and their cargoes to be much simpler. Instead of having to come up with a "last mile" solution for every new passenger or cargo spacecraft, you can have a standardized tug interface, and have the tug do all the hard work. That means that it becomes easier for launch providers to get involved in station resupply, because they're now just taking a standardized container, launching it to a specific orbit, and holding attitude until the tug can swing by and pick things up. Right now, most crew or cargo deliveries to the station require a system that uses a complicated service module and prox-ops hardware to actually get to the station, which results in fairly poor launched mass to delivered mass ratios. What tugs allow you to do in the cargo case is to drastically reduce the amount of wasted mass required to deliver a given mass of cargo to a station. Instead of having your cargo vehicle be a fully capable spacecraft, all it is now is a pressure shell, with some tug interface attachment (probably something brutally simple involving a couple of "hand holds"), and a passive CBM adapter on the other end. If you're launching to a station that's in a resonant orbit that provides frequent "first or second orbit rendezvous" opportunities, you might even be able to dispense with the need for power, communications, or even much in the way of thermal management. In other words, the cargo container starts becoming a lot more like your dumb intermodal container that you see on earth (just much lighter...). Tugs can also serve an emergency role for spacecraft that do have their own prox-ops capabilities, by serving as a backup in case something breaks (or in case multiple docking attempts need to be made and the visiting vehicle runs out of maneuvering propellant). Tugs are also a critical enabler for propellant depots. For propellant deliveries, the propellant can go through relatively narrow tubes (compared to what a human could fit through for instance), which means that a tug could allow for a very simple and lightweight standardized propellant transfer interface to be developed that could just be welded into the delivery tank. This interface could be 100% passive--just some mechanical attachment points, and the quick disconnect ports for fluid and if necessary power. A tug with robotic arms could then take all of the complexity onto itself for the fluid coupling. Much better than trying to make an automated docking and fluid coupling system that has to fly on each and every propellant delivery.

In a nutshell, tugs allow you to take all of the most complicated parts of getting people, propellants, and provisions to a station, and offloads it to either the launch vehicle, or to a reusable vehicle that always stays in orbit, doesn't have to reenter, etc. Why lug all of that hardware with you each and every time if you can leave it at the destination. Why require each and every company that wants to launch stuff to a station to then also have to come up with their own prox-ops solution? Solve the problem once, and then you don't have to keep solving it again. If your delivered payloads start outgrowing your tug, the right option might be to build more of them and operate them in a group, instead of designing a newer, bigger model. I think tugboats do just that for very large ships here on earth--instead of building a super jumbo tug, they'll often just use two or three smaller ones.

Ferries provide very different benefits. First off, and most importantly in my opinion is the fact that ferries (when combined with propellant refueling capabilities) allow you to launch a given exo-LEO vehicle on a much smaller, higher-flight rate vehicle. Dave Salt has on many occasions mentioned that an RLV with an 8000-9000lb payload capability could pretty much service the entire GEO satellite market. Most of the mass required in LEO (I know, many GTO launchers don't even stop in LEO, but it's still a useful point) to put a satellite into GEO is not the satellite, or its "beginning of life" propellants--it's the upper stage, its propellants, and the circularization propellants on the satellite. By having a ferry that operates between LEO and GEO, that has refueling capabilities in LEO, you can launch the largest commercial and government exo-LEO missions without requiring anything bigger than a bottom-of-the-line EELV. In fact, you can even launch manned lunar missions using launchers no bigger than an Atlas V 401 or a Falcon IX (a "Phase One" Atlas V might be a little nicer, but not because of the extra payload to LEO, but because the ICES stages envisioned are scalable and potentially much bigger than a stock Centaur stage, and would thus make a great starting point for a passenger transport ferry). For geostationary satellites, ferries can provide an extra service. Because the ferry can deliver things all the way to GEO, the satellite they're carrying could possibly forgo its "main propulsion system" and circularization propellant tanks in exchange for more station keeping tanks, more transponders, more solar panels or what have you. Or, you could leave the main propulsion system on, but have the capability to retire the satellite to a different, lower-value GEO slot, where it could spend its last few years before moving itself to a final disposal orbit. For instance, by the time a satellite is nearing 15 years on orbit, it may be a bit obsolete for first-world markets, but maybe it would still be useful for a different GEO slot servicing locations in the third world, or sparsely populated areas in the Pacific for instance (much like how passenger jets in the US are often "retired" only to be refurbished a bit and sold to third world countries at a much lower price). Either of these can help you get more revenue out of a given satellite launch. There are probably plenty of other benefits of ferries that I'm not thinking of right now, but those are just some thoughts.

Ferries can be based around either chemical or solar electric propulsion systems. Some cargoes don't mind a slow spiral out through the van Allen belts, and thus can be shipped by the more mass efficient (and hopefully therefore more cost efficient) solar-electric "slow boat". Other cargoes (people, cryogens, and possibly GEO satellites) can be shipped via a much faster chemical ferry. Sure, it's less mass efficient, so you're going to be paying for launching a lot more material, but the hardware is relatively cheaper, it can make more flights before being retired, and most importantly, you're not cooking your payload for several weeks in the van Allen belts. For GEO satellites right now, most of their radiation exposure (for their entire 15 year operation timeframe) happens in just one or two passes through the van Allen belts, so minimizing the time spent there might give chemical ferries a leg up (contra conventional wisdom).

Anyhow, what do you guys think? Does drawing this distinction make sense? And does anyone have a term better than "ferry" for a reusable transfer vehicle? Every time I've tried to bring up the idea of a "space ferry" there at the conference, the term would draw smirks or chuckles, or comments along the lines of "I guess NASA Ames is close to San Francisco after all"...

Orbital Access Methodologies Part II: The Key Challenge of TSTO RLVs

Before I go into detail on any of the two stage to orbit (TSTO for the uninitiated) approaches that I mentioned in my post last week, I'd like to briefly discuss what I think is the key issue that drives the design and development tradeoffs for reusable TSTO launch vehicles. That issue is: how do you get the first stage back after a mission, and ready to fly again?

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.