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.

Welcome to the Club!

Zond, on NASASpaceflight.com, just posted a link to a new European commercial space venture, Orbspace. Like Masten Space Systems and Armadillo Aerospace, Orbspace is pursuing a suborbital VTVL vehicle, which they intend to use for both manned and unmanned applications. As I said over there on NASASpaceflight, I think this a really positive development. While there have been several international companies that have announced ambitions to field suborbital vehicles, only some of them have been taking an approach that I thought gave them a good chance to realistically go anywhere. One of the others that I think is taking a decent approach is Project Enterprise (some of the people involved in that project are friends of mine), which is partnered with the Swiss Propulsion Lab. If Orbspace is trying to be a European equivalent of MSS or Armadillo, Project Enterprise is more like an attempt at a European XCOR.

There are some bloggers who seem to be perpetually stuck in a Cold War mentality, always looking for the next USSR, and a return of the glory days of the Space Race. The fact that this would be totally counterproductive is apparently lost on these people (as is the weirdness of people who claim to be free-market capitalists rooting for socialist design bureaus that would've been right at home in the former Soviet Union). I really think that it is entrepreneurial space ventures like Orbspace and Project Enterprise that really represent the future of international space competition. I'm not afraid at all of the Chinese national space program beating us back to the Moon--they're following a dying and deprecated model of space development that hopefully won't last too much longer into this century. What does cost me a little sleep every now and then is wondering what's going to happen when Chinese entrepreneurs start seeing the success of US and European commercial space groups, and decide that they could make a buck at that too...

Welcome to the club guys!

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"...

It's Not Important Whether You Win or Lose...

...it's how you place the blame. Or at least that's how a friend of mine at NASA once put it.

Apparently, in the wake of yet more news leaks about severe technical issues on Ares I, Mike Griffin decides to play the blame game (hat tip: Space Politics):

A: Let me get down to the bottom of it. There were winners and losers in the contractor community as to who was going to get to do what on the next system post shuttle. And we didn’t pick (Lockheed Martin’s) Atlas 5, in consultation with the Air Force for that matter, because it wasn’t the right vehicle for the lunar job. Obviously, we did pick others. So people who didn’t get picked see an opportunity to throw the issue into controversy and maybe have it come out their way.

I'm sure the guys in Denver are getting a hoot out of this. Of all the people who have a reason to be upset at the massive waste that is Ares-I/Ares-V, the LM/ULA guys I know have actually been rather politic about their complaints.

You see, they're too busy working on making a vehicle that's safe enough to fly people and affordable enough to do so entirely on the private market, without needing multi-billion dollar sole-source contracts from NASA. Of course, Mike may actually have a point though. If there is anyone outside of NASA HQ who is to blame for making Ares-I look bad, it probably is the Atlas V guys...

Just not in the way Griffin is insinuating...

LM/Bigelow Atlas V Deal

For those who didn't see it on Hobbyspace, I got interviewed yesterday by New Scientist about the recent LM/Bigelow announcement. All in all it was a pretty good article (though apparently we might need to update our website to reflect the fact that we haven't been in Santa Clara for over a year and a half...). I had a few other thoughts about the announcement that I figured might be worth sharing, for what its worth.

In the quote they selected for the article, they mentioned my question of "will they be able to drum up enough demand to justify the flight rates they're talking about." Here were some of my thoughts that I shared with David Riga (the author of the New Scientist piece), that didn't make the cut:

If he were just running an orbital hotel (he isn't), I'd be very skeptical. Instead I'm somewhere between skeptical and guardedly optimistic. While there haven't been large numbers of takers for flights on the Soyuz, what Bigelow's offering is fundamentally different. Flight opportunities are frequent (which is critical for most microgravity research programs--imagine trying to run an R&D lab that you could only visit once or twice a year!), the situation is more customer friendly, training would likely be more streamlined (I hear that for Soyuz training the "passenger" is actually more of a third crew member than an honest-to-goodness passenger), etc.

It'll be interesting to see if he can pull off his idea of forming an international astronaut corps for countries that don't have their own space program. It wouldn't have all the usual glory of having your own national launch system, but it also wouldn't have the waste of it either. Countries like the UK could look at it as a smart and low-cost way of doing a manned space program--why reinvent the wheel when you can just buy a ticket and focus on doing something in space instead of blowing billions just getting there?

Also, the title of the New Scientist piece is somewhat misleading (though David may not have had anything to do with the title). There are some major hurdles for using Atlas V to fly people to Bigelow's station--it's just that most of the major risks don't lie with "man-rating" the Atlas V (or whatever you want to call making reasonable adaptations for flying a capsule on an ELV). Continuing with some more thoughts that didn't make the cut (yeah I wasn't expecting David to use every word of my several page response...):

Most of the challenges fall into two areas: developing a market at the pricepoint Bigelow can offer with existing transportation systems (like a "man rated" Atlas V), and finding a capsule developer who can raise the money and technically execute on doing such a capsule. I think the technical risk for both parts is relatively low--this has been done before even if there are still some improvements needed over previous systems (Mercury, Gemini, Apollo, Soyuz, etc) to make it commercially viable. Most of the risk is on the marketing and financing side of things.

If Bigelow is able to start signing up high-visibility customers though, look to see that change. Once there looks like there's going to be enough demand to justify a capsule project, I think it'll be much easier to raise money for [developing] it.

Lastly, discussing whether I thought that the Atlas V was a good choice for Bigelow, I said:

I think at the moment they're a pretty good choice. The good news is that with SpaceX also hopefully getting into the launch business soon, that'll provide the competition Bigelow needs to keep prices low. Obviously, it would be great if there were high-flight-rate commercial RLVs instead, but those really need a proven market in order to justify the funds needed to pull them off. So short term, I think this may be Bigelow's best bet. In the longer term, it'll be up to LM to find ways
to keep themselves competitive.

To elaborate on this last point a bit, the price points Bigelow has been talking about (~$15M per person for a 1 month stay) and which a system based off of the existing Atlas V could likely deliver are probably too high for there to be a lot of space tourism demand. Fortunately, as Bigelow has mentioned a lot of times, he isn't running a space hotel. In order to really start getting to the elastic portion of the demand curve, the price tag would probably need to be a bit lower--on the order of $2-5M per ticket (according to some reanalysis of the old Futron Space Tourism study that T/Space did a few years ago that I discussed in this old blog post). It may not actually be as impossible for LM to deliver numbers at least on the high-end of that scale as I used to think (they have some possible tricks up their sleeve if the demand for Atlas V flights was high enough to justify the investment), and if Bigelow can actually deliver on demand for 80+ people to his station in a given year it might also be enough to close the business case for a high-flight rate, small RLV. But neither of those options are likely to happen right away. So, while someone like Space Adventures could probably rent some of his facility for space tourists, at the price point they are talking about, I'd be surprised if they could fill up more than 1-2 of the 12 targeted flights per year with actual "space tourists".

That leaves Bigelow's "sovereign" and "prime" customers to make up the rest of the 10 flights worth of demand. Admittedly one should note that not all of the 12 flights per year are going to be people--I'd imagine that at least one will be consumables, cargo, reboost propellants, etc. And on some flights I imagine that some of the passenger seats might be exchanged for experiments, research hardware/raw materials, and other commercial cargo.

The good news is that if they're really providing 12 missions per year, that's a monthly flight. While that still isn't phenomenally great for a microgravity research program (see Ken's last post, and my last space post and these posts from the ACES conference two years back for why flight rate is important for such programs), it's substantially better than the existing state of practice. As was stated in the first of those two ACES posts, when people know that there's going to be a flight every month to the station, it's a lot easier to slip last minute experiments or small hardware on-board at the last minute. Scientific research often lives or dies on iterations--on how fast you can experiment, analyze, reformulate, rehypothesize, and get to your next experimental step. What this means is that while 12 flights a year at $15M per seat isn't perfect for orbital microgravity research, it might actually be good enough to start generating some real demand--ie the "tipping point" where orbital microgravity demand starts picking up might be a little higher than orbital tourism, and possibly high enough to fill up at least a chunk of those 10 remaining flights.

But like the space tourism demand, that demand is only going to be able to grow if Bigelow can provide enough demand for the rest of those flights. Which brings us back to the "sovereign" customers that Bigelow has mentioned on several occasions. The idea being that this would provide smaller countries a much cheaper way to get involved in manned space flight. At least one country I know of might be in an ideal position to take the lead on this venture: the UK.

As Duncan over at the Rocketeer blog has mentioned on several occasions, this might be a good way for the UK to get back into manned spaceflight as they have recently been discussing more seriously. It's interesting to note that the premier suborbital tourism venture involves a US launch provider and a British operator, so the idea of the UK buying tickets to a US owned commercial station on US owned and operated launch vehicles could be framed as being the new way of doing things. As I mentioned above, by letting someone else spend the money on the destination and the transportation, the UK could focus on actually doing something useful with people in space, instead of blowing so much money on the first two categories that they have little left for actually accomplishing something. This would be a very forward-thinking thing for the UK to do. And if they took the lead in signing up for such a program, it is very feasible to believe that you would see other nations following their lead. I'm thinking of other Anglosphere countries like Canada, Australia, New Zealand, South Africa, and possibly even India. It wouldn't take too many of them running small low-cost astronaut corps and doing their own research projects on Bigelow stations before you could start providing enough demand to see those kinds of flight rates. Or at least it doesn't seem to unrealistic to imagine it.

So, at least on the surface it might be possible for Bigelow to pull this off--but he's going to need to sign up some high profile customers sooner rather than later. In the medium and long term, if Bigelow is able to provide enough demand for that many Atlas V flights, LM is going to have a lot of competition. From SpaceX and from other corners. But that's a problem that I'm sure we would all love to have...

Discussion of Dr. Griffin's STA Comments on ESAS

I've had several people in several places ask me if I was going to do a point-by-point rebuttal of Mike Griffin's comments to the STA this week (for reference the text of his comments is available here). While I don't have the time to go into every single disagreement I have with what he said, I think there are a couple of key points I would like to point out. In other words, I've come to discuss Griffin, not to Fisk him.

Missing the Vision

Dr Griffin starts his defense of the chosen Constellation architecture by framing it "in the
context of policy and law that dictate NASA’s missions." As he said on page 2:

Any system architecture must be evaluated first against the tasks which it is
supposed to accomplish. Only afterwards can we consider whether it accomplishes
them efficiently, or presents other advantages which distinguish it from competing
choices.

He then went on to discuss President Bush's original announcement of the Vision for Space Exploration, and the NASA Authorization Act of 2005. I agree that it is important to make sure you know up-front what yardstick your program is going to be measured by. However, I think one thing becomes quickly obvious as you read Dr Griffin's quotes from those documents--he entirely focuses on the technical implementation details, and never once mentions the actual policy goals!

Quoting from "A Renewed Spirit of Discovery: The President’s Vision for U.S. Space Exploration":

Goal and Objectives
The fundamental goal of this vision is to advance U.S. scientific, security, and economic interests through a robust space exploration program.

These goals are the yardstick by which any VSE implementation needs to be judged. The rest of the technical details of how the space exploration program is carried out needs to be viewed in the light of these three areas of US interests. It doesn't matter if a proposed implementation hits all of the other technical details, if it doesn't really further US scientific, security, and economic interests, it isn't really compliant with the goals of the president's Vision.

Going into a little more detail on these goals, the Renewed Spirit of Discovery document continues (emphasis mine):

In support of this goal, the United States will:
• Implement a sustained and affordable human and robotic program to explore the solar system and beyond;
• Extend human presence across the solar system, starting with a human return to the Moon by the year 2020, in preparation for human exploration of Mars and other destinations;
• Develop the innovative technologies, knowledge, and infrastructures both to explore and to support decisions about the destinations for human exploration; and
• Promote international and commercial participation in exploration to further U.S. scientific, security, and economic interests.

Once again, all of the specific technical details like the CEV, retiring Shuttle in 2010, etc. are all pursuant to these goals.

Lastly, the NASA Authorization Act of 2005 (available here) states, once again with my emphasis:

The Administrator shall establish a program to develop a sustained human presence on the Moon, including a robust precursor program, to promote exploration, science, commerce, and United States preeminence in space, and as a stepping-stone to future exploration of Mars and other destinations.

Once again, you will notice that the key goals of this Vision, elucidated by both the President and Congress include not only science, but commerce, and in the president's case security.

I could go on about how Dr Griffin's focus on the parts of the Authorization Act that talk about heavy lift and shuttle derived ignored other sections in the act that talk about "encouraging the commercial use and development of space to the greatest extent practicable" (see Section 101.a.2. parts B-C). But I think the fundamental issue is that by focusing exclusively on just the technical side of the requirements, and not on the underlying goals, Griffin is missing the Vision.

Growth Potential

On page 7, Dr. Griffin starts making his case for the Constellation architecture with this somewhat ironic statement about the Space Shuttle:

Once before, an earlier generation of U.S. policymakers approved a spaceflight architecture intended to optimize access to LEO. It was expected – or maybe “hoped” is the better word – that, with this capability in hand, the tools to resume deep space exploration would follow. It didn’t happen, and with the funding which has been allocated to the U.S. civil space program since the late 1960s, it cannot happen. Even though from an engineering perspective it would be highly desirable to have transportation systems separately optimized for LEO and deep space, NASA’s budget will not support it. We get one system; it must be capable of serving in multiple roles, and it must be designed for the more difficult of those roles from the outset.

And then Dr Griffin goes on to try and justify an architecture based on building a duplicative LEO capable only launch vehicle first, and hoping that when that vehicle is finally done, that there will be funding for developing "the tools to resume deep space exploration"...

After that auspicious start, Dr. Griffin then reminds us that "the new system will and should be in use for many decades." Of course some of the historical analogies he draws could lead one to different solutions than it led him. For instance, he mentions that "In space, derivatives of Atlas and Delta and Soyuz are flying a half-century and more after their initial development." An interesting thing to note about Atlas and Delta is that the only reason why vehicles with the name Atlas and Delta are "still flying" a half-century after their initial development, is precisely because they are only derivatives of the original. In fact, the current EELVs have very little in common with the vehicles that originally bore their names.

On pages 8 and 9, Dr. Griffin concludes that (emphasis mine):

The implications of this are profound. We are designing today the systems that our grandchildren will use as building blocks, not just for lunar return, but for missions to Mars, to the near-Earth asteroids, to service great observatories at Sun-Earth L1, and for other purposes we have not yet even considered. We need a system with inherent capability for growth.

While I disagree with the direction Dr. Griffin is going, I do agree with his point in that last sentence. We do need a transportation architecture that has inherent capability for growth. I just don't think that the Constellation architecture really fits that bill.

The Promise of Commercial Space

Now, lest you think I'm going to spend yet another post hammering on Dr. Griffin, I'd like to quote a part of his speech that I really agreed with:

Further application of common sense also requires us to acknowledge that now is the time, this is the juncture, and we are the people to make provisions for the contributions of the commercial space sector to our nation’s overall space enterprise. The development and exploitation of space has, so far, been accomplished in a fashion that can be described as “all government, all the time”. That’s not the way the American frontier was developed, it’s not the way this nation developed aviation, it’s not the way the rest of our economy works, and it ought not to be good enough for space, either. So, proactively and as a matter of deliberate policy, we need to make provisions for the first step on the stairway to space to be occupied by commercial entrepreneurs – whether they reside in big companies or small ones.

I have to say that for all my disagreements with Griffin, he at least talks a good talk when it comes to commercial space. I full-heartedly agree with his point in this paragraph. When you think about it, even assuming everything works out according to his plan, Constellation is never going to be capable of supporting more than a dozen people off-planet at any time. While that may be a lot more than we have now, Ed Wright has a point when he says that that is a round-off error, not an exploration program. Basically, the only way we're going to see large numbers of people off planet, and the only way we're going to see the large-scale manned exploration and settlement of our solar system in our life times, is if the private sector can eventually play a much more expansive role in space transportation. As it is right now, so long as the commercial industry continues to play second fiddle to parochial interests and NASA-centricism, we're not really going to go much of anywhere.

So, the fact that NASA is at least doing something to help promote that day is a sign that they at least partially get it. A successful and thriving entrepreneurial space transportation industry is going to help them actually achieve their goal of extending human life throughout the solar system in a robust program of space exploration.

Griffin continues with more good comments in his next paragraph:

If designed for the Moon, the use of the CEV in LEO will inevitably be more expensive than a system designed for the much easier requirement of LEO access and no more. This lesser requirement is one that, in my judgment, can be met today by a bold commercial developer, operating without the close oversight of the U.S. government, with the goal of offering transportation for cargo and crew to LEO on a fee-for-service basis.

But here is where the conversation takes a dangerous turn:

Now again, common sense dictates that we cannot hold the ISS hostage to fortune; we cannot gamble the fate of a multi-tens-of-billions-of-dollar facility on the success of a commercial operation, so the CEV must be able to operate efficiently in LEO if necessary. But we can create a clear financial incentive for commercial success, based on the financial disincentive of using government transportation to LEO at what will be an inherently higher price.

To this end, as I have noted many times, we must be willing to defer the use of government systems in favor of commercial services, as and when they reach maturity. When commercial capability comes on line, we will reduce the level of our own LEO operations with Ares/Orion to that which is minimally necessary to preserve capability, and to qualify the system for lunar flight.

While I agree that the government not only is the government being "willing to defer in favor of commercial services" is a really good idea, I think that this approach (of hedging their bets by coming up with a competing in-house launcher) is fraught with risk. Also, while on first blush, it may appear to be common sense to not "hold the ISS hostage to fortune", it is my contention that this line of reasoning not only doesn't hold as much water as it seems.

First off, as has been pointed out on numerous occasions, including in Griffin's statements above, a commercial solution to ISS crew/cargo is going to be a lot more affordable than the in-house Ares-1/Orion solution. It has been mentioned before by people high up at NASA, that they really need COTS to succeed, because if they have to fly all the ISS missions themselves (especially if ISS doesn't get retired in 2016, which Dr. Griffin mentioned in this speech as a possibility), there really won't be anywhere near enough money to develop the lunar portions of the proposed Constellation architecture in time for the 2020 lunar return goal. You could say in a way that the existing Constellation architecture holds the rest of the Vision hostage to the fortune of COTS. If COTS doesn't succeed, there's no way NASA is going to be able to afford executing on the rest of the vision. If the supposed "backup plan" for ISS resupply won't produce acceptable results anyway if COTS doesn't turn out, NASA shouldn't be trying to make it a backup plan at all--they should invest more heavily in making sure that there are multiple COTS competitors and that they have enough resources to succeed. One of the single biggest execution risks for any COTS company is financing risks. And having a NASA "backup plan" that could potentially compete with them is one of the single biggest obstacles to be overcome in raising money for a COTS team.

Which brings me to my other concern. The danger of having NASA in-house launch vehicles and space access capabilities that can serve as a backup to COTS also allows them to directly compete with COTS if the budgetary situation goes sour. Think about it. If Ares-1 finally gets built and working, but Ares-V doesn't get funded, there's nothing for Ares-1 to do but service ISS. With how hard the esteemed congressmen from Florida, Utah, and Alabama are fighting to maintain the Shuttle workforce and infrastructure (even to the point of suggesting continuing to fly the Shuttle!), does anyone really think that they would just "stand down" at that point, even if there was a clearly superior commercial alternative? Not very likely. I'm sure they would come up with some technical reason why Ares-I was superior (after all, our probabilistic risk assessment says that Ares-I has a 1:2106.5923 chance of killing a crew, while our numbers show that they have a 1:500 chance--who do you want flying our brave astronauts?) and find a way to not actually stand down. The frustrating thing is that by setting things up the way NASA is doing, the NASA people don't even have to be malicious for such a result to happen--it's a natural and likely consequence of the perverse incentives that NASA and Congress are setting up.

So, while I personally think that Dr. Griffin really and emphatically believes in and supports commercial space development, I'm afraid that there's a high chance that some of his well-intended choices could end up coming back to haunt us.

Moon, MARS!!!! and Beyond

The last item I'd like to point out in Dr. Griffin's speech is one of the justifications he used for the "1.5 launch" architecture they selected. Dr. Griffin made the point that while he feels that Constellation needs to be backward compatible with ISS as a backup plan, it also needs to be forward compatible with Mars, because sometime in the 2030s, we're going to be going there. Now, I'm of the opinion that trying to guess what the best technical approach will be for a problem 30 years from now is somewhat of a fools errand. But that's just me I guess.

So, starting on page 16 he begins to layout his case:

On the other end of the scale, we must judge any proposed architecture against the requirements for Mars. We aren’t going there now, but one day we will, and it will be within the expected operating lifetime of the system we are designing today. We know already that, when we go, we are going to need a Mars ship with a LEO mass equivalent of about a million pounds, give or take a bit. I’m trying for one-significant-digit accuracy here, but think “Space Station”, in terms of mass.

Now, I'm not going to go into the fact that there are probably plenty of other approaches to Mars exploration that can change the equation entirely. That's a post for another day. For now, let's just run with that premise.

He then repeats the "everyone knows that ISS taught us that using 20 ton vehicles to build something big is a bad idea" catechism, but that's not what I'd like to discuss. The real gem is in this paragraph on page 17 (emphasis mine):

But if we split the EOR lunar architecture into two equal but smaller vehicles, we will need ten or more launches to obtain the same Mars-bound payload in LEO, and that is without assuming any loss of packaging efficiency for the launch of smaller payloads. When we consider that maybe half the Mars mission mass in LEO is liquid hydrogen, and if we understand that the control of hydrogen boiloff in space is one of the key limiting technologies for deep space exploration, the need to conduct fewer rather than more launches to LEO for early Mars missions becomes glaringly apparent.

It is true that one can draw that inference--that hydrogen boiloff means you should build as big of an HLV as possible. However, the conclusion I would draw is that if cryogenic propellant storage technologies are "key limiting technologies for deep space explortion", then the right answer is to stop trying to kludge around the problem--develop them! Don't use the existing state of the art in propellant handling and problems that are still 20 years down the road drive multi-billion dollar development projects today.

There are current technologies under development that could yield very low to zero boiloff of cryogenic propellants. There are multiple groups (ULA, Boeing, groups working with Glenn Research Center, etc.) pursuing multiple approaches to solving these problems. There are passive cooling and active cooling techniques. This isn't some high-risk technology like nuclear fusion. The technologies needed for cryogenic fluid management in space are mostly low-risk extensions of 40 years worth of research and development. More to the point, many if not all of these technologies need to be developed to make Constellation work for lunar trips anyway, and would still be needed for Mars trips.

Is 2030 really so close that we can't afford to do this right and actually develop the technologies we need instead of trying to kludge by with existing technologies?

Once you have the boiloff issue reduced or solved, that ~500klb of hydrogen ceases to be a headache, and begins to be an opportunity. That's a lot of demand for propellant in orbit, and it can be supplied commercially. You're already going to need propellant transfer technologies anyway if you have to launch the hydrogen in multiple launches, so what's to stop launching it in even smaller launches?

I guess my point is that if one of the key arguments for the 1.5 launch architecture over a more commercial one, or a less expensive shuttle derived one like DIRECT is hydrogen boiloff, I think their kludge around the issue isn't the right approach, and that they'd be better off just doing it the right way. Also, part of the reason why we have a federally funded aerospace program is to help prove out the technologies necessary for enabling the commercial exploitation of space, and actually solving problems like these would be much a much more responsible use of public funds than developing a kludge around point design like Ares V that doesn't advance the state of the art for the commercial benefit of the country.

Conclusions

I guess overall while there were some good points, there was also a lot of issues with Dr. Griffin's latest defense of Constellation. As discussed, I think that an a myopic focus on the technical details while ignoring the overall goals of the VSE has led to an architecture that isn't responsive to the key policy goals laid out by the president and reiterated by Congress (particularly with respect to promoting the commercial and security interests of the United States). I think that in spite of Griffin recognizing the need for growth and flexibility in any architecture, that he chose a rather brittle and inflexible one. I also think that while he showed that he does recognize the potential of commercial space, and the importance of NASA trying to promote it, I think that the way he's running COTS and Constellation will likely end up being highly counterproductive. Lastly, I think that in many cases, when confronted with a solvable engineering problem, Constellation has instead decided to kludge around the problem instead of properly solving it.

There are plenty of other issues I could've raised, but I figured these were some of the more obvious ones that I felt needed discussion.

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 th