Technology Lock-in
One of the most interesting points made in the paper was how often technology can get locked-in by early decisions in a field, often made in situations of very limited data.  As documented in the paper, the engineers developing the Apollo docking  system admitted that “The selection of a docking system for the Apollo Program was based on limited knowledge because experience with actual hardware in space or from ground-based docking simulations was almost nonexistent.”  These early conceptual downselects, in this case with only a few half-scale air-bearing experiments to provide any data at all, often are not revisited in future developments.  This puts us in a situation where hasty decisions made early-on, based more on “engineering judgment” end up getting stuck with for many decades after the fact.   While this paper was focused on rendezvous and docking techniques, there are plenty of other examples of similar technology lock-ins in aerospace and elsewhere in industry.

Illustration of Boom Rendezvous
I found the following illustration, pulled from a NASA presentation given by Kirk Sorensen of MSFC (one of Joseph’s coconspirators on many topics including air launch, MXER tethers, and Thorium Liquid Flouride Reactors), to be the best illustration of the concept:

Much like modern mid-air refueling of helicopters and jets, a low-inertia connection is made at the end of booms extended from both vehicles, instead of trying to actively fly the two vehicles into each other.   In this case, the boom is on the order of 10-100m long, giving plenty of space to avoid collisions while hooking up the booms.

Boom Design Concept

The preferred approach for the extendable boom was to use a system like the Bi-STEM, which has been manufactured by Northrop-Grumman for in-space applications for decades:

The Bi-STEM system is sort of like a pair of tape-measures.  The coils of spring-steel form into arcs as they are spooled out, and their ends are interlocked, creating a tube that can be actively lengthened and shortened using electric motors or the spools.  A polymer tether can be easily run down the center of the assembly, adding greatly to the system’s tensile strength:

The whole assembly can be mounted on a gimbal or a Camfield Joint allowing full 6DOF pointing, using electric actuators.

Advantages

The main advantages of Boom Rendezvous, as detailed in the paper linked to, include:

• Greatly decreased probability of collisions during rendezvous.  Of all the possible rendezvous failures, collisions are the most likely to badly damage one or both vehicles.  Being able to greatly reduce the probability of damage due to failed docking is critical for operations like propellant depots that may require hundreds or thousands of successful docking operations over their lifetime.  With Boom Rendezvous, a missed connection goes from being a serious hazard to being the kind of thing you can easily try again, with the only risk being the ego risk of getting razzed by your fellow astronauts after the fact.
• Greatly reduced propellant requirements for the docking maneuvers.  Instead of the hunting problem often faced in current real-world docking operations, the closing is performed almost entirely by electric motors.
• Elimination of plume impingement problems.  When maneuvering two rocket-powered vehicles close together, impingement of the jets from the maneuvering vehicle on the other vehicle structures can be a severe problem.   Impinging plumes can spall off structural material or contaminate surfaces and optics.  Since all the maneuvering close-in is done using the booms, this is eliminated.
• Much lighter and simpler connection interfaces, since the booms can eliminate any remaining rotational or angular misalignments, and since the booms up close have enough compressive strength that you can very precisely control the final connection loads.  Without those extra loads, you can eliminate the heavy backing plates, shock absorbers, and guide petals common in modern docking adapters.  And without having to have those in the middle, the latching mechanisms, seals, and fluid/electrical connections can be made a lot more straightforwardly.
• Reduced sensor requirements for the rendezvous/docking.  You no longer need to know anywhere near as much about the target vehicle’s velocity and orbit, which allows you to use less sensitive, more robust sensors to make the hookup.
• Less precision required in the initial rendezvous orbit.  This may allow for the upper stage of the launch vehicle to do the rendezvous burn, allowing the payload to be much simpler and “dumber” than your typical modern prox-ops stage (like Dragon or Cygnus).  This will be important for depot operations as well, because the less smarts the tanker has to have, the lower it’s cost can be, and the more space can be left for the actual useful cargo or propellant.

Applications of Boom Rendezvous
While Boom Rendezvous has many benefits compared to the existing probe-and-drogue based docking systems in-use today, it has a few areas where it really shines:

• Rapid Rendezvous situations.  These include MXER tethers, apogee tugs, exo-atmospheric suborbital refueling, and other situations where the vehicle needs to hook-up very quickly.  Trying to do that with the high-inertial close-in maneuvering typical of today’s rendezvous and docking systems is begging for a crash.
• Depots and other space facilities.  The ability to have the actual docking with a depot occur several meters away from sunshields, tanks, and other hardware increases the odds of the depot being able to last long enough to be economically useful.  In fact, it may be possible using Boom Rendezvous for the Tanker/Tankee to offload or onload propellants without ever actually touching the depot itself.
• RLVs.  Most early-generation RLVs are likely going to be rather weight constrained.  By providing a potentially lighter docking system that doesn’t require as many demanding subsystems, more weight can be reserved for payload and recovery systems.
• Space Tugs.  Boom Rendezvous makes it a lot easier to divide the docking up into as many tugs as are necessary for safe operations.  Much like how more than one tug boat can be used when bringing a large oceanliner in to dock.

Concluding Thoughts
Boom Rendezvous and Docking is a rather promising approach that I hope sees more investigation.  The cool thing is that with the advent of suborbital vehicles, this is the kind of system that could be rapidly matured and demonstrated “in-space” for a tiny fraction of what it would cost to do with purely orbital systems.  Hopefully, the changing technological maturation situation provided by reusable suborbital launch vehicles can allow us to finally revisit hasty decisions made during Apollo.

1. Parallel Ascent Air-Launch Glide-forward TSTO with Exo-Atmospheric Refueling.  A crazy idea I’ve been stewing on involving the intersection between air-launched TSTOs, propellant transfer, FLOC, and Boom Rendezvous, and my attempt at making it seem a little less crazy than it sounds.  If it can be made to work, it might allow for fully reusable manned orbital launch off of carrier planes as small as WK2.
2. Boom Rendezvous.  An introduction to a clever rendezvous and docking technique that I think deserves more investigation.
3. Dual Fluid Propellant Depots.  A little more in-depth discussion of my favorite depot concept from my prop depot paper.
4. Refuelable Un-Crasher Stages for Early Lunar Lander Reuse. A discussion of a lunar lander architecture I came up with that leverages propellant depots to enable early reuse of a significant part of the lunar lander system, while still allowing impressive (ie >>ESAS) landing capabilities.
5. The Slings and Arrows of Outrageous Lunar Transportation Schemes.  A discussion of several of the non-rocket methods for getting payloads off of (and in some cases back onto) the Moon.  Including pro’s, con’s, and technology hurdles.

Not too much yet on the why’s, but I’ve been keeping these how’s bottled up for too long.  Which of these sounds most interesting to you guys?  In the comments, pick your favorite one or two, and say why.

I’ve been so busy lately that I’m only now getting a few spare minutes to show off some rather cool pictures I was sent a week ago.  A week ago Monday, I got an email from Claire Flanagan of the Johannesburg Planetarium in South Africa, asking about some strange pictures, including the one taken above by Mitchell Krog of Livingcanvas.co.za (check out some of his nature photographs–they’re breathtaking!). She had reason to believe that these bubbles had something to do with last week’s Atlas-V DMSP launch.

This was the launch I mentioned in my propellant depot paper, where after delivering the spacecraft they did a series of cryo-fluid management tests, including several tests on centrifugal propellant settling. BTW, I hear they got some good data from the tests, but I’ll let the ULA guys speak for themselves once the data is fully analyzed.

Anyhow, after helping the Johannesburg Planetarium gals/guys get in contact with ULA, they were (with Case Rijsdijk’s help) able to resolve the mystery. But I’ll let you read what Case and the Planetarium had to say about it, instead of just repeating it.

Kudos to Mitchell and Claire for letting me post some of the pictures and links.

First, the propellant depot paper I keep talking about: AIAA 2009-6756 Near-Term Propellant Depots: Implementation of a Critical Spacefaring Technology (also available on the ULA publications page), and the presentation that went along with it: SPACE 2009 Prop Depot Paper Presentation.

For those who’ve read most of the rest of my posts on propellant depots, there’s only a few key new concepts:

• The Single-Launch Dual-Fluid Depot concept that Frank Zegler and myself both independently came up with this year (which I’ll go into more in a later blog post).  This idea holds a lot of promise because it shows how a single EELV launch with existing fairing sizes can put up a depot capable of storing 75-114mT of LOX/LH2.
• The realization that unless you have some sort of high-Isp or propellantless propulsion system (like an ED tether) for stationkeeping purposes, that Zero-Boiloff storage might not be very useful for an LEO depot.  The amount of propellant you lose to boiloff is less than the amount you would’ve spent for stationkeeping anyway, so it’s effectively free.
• The realization that LEO depots really ought to be treated as “use-it-or-lose-it”, high-throughput depots, and that its L1/L2 depots that should be used for longer-term storage.

I had been avoiding discussion of the dual-fluid depot concept for a while, mostly so I wouldn’t be stealing my own thunder.  Now that the paper is presented and out in the public, I hope to have the time soon to discuss the concept a bit.  I also have some space transportation architecture ideas using that depot concept that I may flesh out a bit either here on the blog or in some white papers (which I’ll post on the blog).

The second paper, which was mostly written by Robert Frampton of Boeing talks about a project to use our XA-0.2 vehicle as a testbed for demonstrating autonomous landing systems for planetary landers: AIAA-2009-6571 Planetary Lander Dynamic Model for GN&C

I’m not really sure what I’m allowed to say about the project that isn’t spelled-out in the paper, but I figured I’d bring it to people’s attention, and hopefully at some point in the future I can discuss things more on the MSS blog.

Depot-Centric Human Spaceflight: Strengthening American Industry, Creating a Robust Beyond-LEO Exploration Program, and Enabling the Commercial Development of Space

Anyhow, Jeff managed in 25 minutes to address human rating, depots, whether or not we need heavy lift, technology maturation and R&T investment, and the need for NASA to find new ways to interact with business. I don’t think he could have hit more of my hot-button issues in 25 minutes if he had tried.

Anyhow, I hear that the HSF committee will have video of today’s proceedings up online soon (possibly tonight) for those who didn’t get up at 5:30am PDT to watch. I’ll comment more later.

Whew! I haven’t had this much hope for this nation’s space program in years!

[Update:  Here's the link to the subgroup's presentation (warning, it's a 14MB powerpoint presentation).  All of it is interesting, but Jeff's part starts on page 76 and goes through page 89.  Chris Chyba's section was the last three pages.]

The mission uses the following tricks to make things work:

• Dual Engine Centaur for this mission is stretched by 50% and includes an “Extended Mission Kit” to allow for it to function for the ~5 days necessary for the mission (normal DEC dry mass is ~5400lb, and the EMK is ~1750lb and includes stuff like extra hydrazine bottles, more batteries, deep space navigation upgrades to avionics, sunshields, etc)
• Command module does a powered lunar swingby to go to L2, thus cutting down on overall dV requirements (~750m/s total required, 335m/s per leg), thus allowing for a much smaller CSM (possibly with the service module integrated into the command module).
• The Stretched Centaur and the Lander break into lunar orbit and descend to the surface instead of continuing to L2.  I’m not positive if this allows you to land anywhere on the lunar surface or not (this is one of the few big questions for this mission mode).  This avoids the extra dV requirements you normally get for stopping everything at L2 first.
• Upper stage performs part of the landing burn (between LOI and the descent burn it provides about 1950m/s out of the total 3050m/s needed for LOI and landing).
• RS-68A Upgraded Delta-IVH. This upgrade is already in engine testing and is badly needed by the DoD, so there’s a good chance this will work out.  Expected payload capacity I’ve heard is 27mT for the system.
• Instead of carrying a second stretched Centaur as a payload on one of the flights, the Atlas V 552 uses the stretched Centaur as its upper stage.  In order to tank up the LH2, it carries an LH2 drop tank between the lander and the command module.  It gets transfered right after reaching orbit, and gets dumped shortly before TLI.

Here are the major components of the system:

• Command Module: This module is based on the Apollo outer mold line, but only carries two people, and enough life support consumables for the mission.  I budgetted 11,000lb dry and 3250lb of propellant for the capsule (not including RCS propellants).  I assumed hypergols for the stage, with a crappy 314s Isp.  The Apollo CM wet mass was 12.8klb, and the SM weighed 54klb wet, 13.5klb dry.  However, most of the SM mass was due to the CSM performing the LOI burn for the Apollo Stack.  About half of the dry mass of the CSM was the huge main engine, and a good chunk of the remaining mass was electrical equipment and the huge tanks for the 40klb of propellant.  With modern materials, electronics, a smaller crew, solar panels instead of fuel cells, and the much lower propulsive requirements for the Command Module in this architecture, I think 11klb is actually pretty conservative for such a system.  For another comparison the latest CEV numbers I’ve heard (which are pretty far out of date) were ~18klb for a four person capsule.
• Stretched Centaur Lunar Transfer/Crasher Stage:  As mentioned above, this is a dual engine centaur using two RL10A-4-2 engines, but with a 50% barrel stretch to the tanks.  The tanks are actually less than 40% of the dry mass of a centaur stage, but you also need more helium for pressurization of the larger stage…assuming that the 50% greater propellant load requires a 50% higher dry mass should be a conservative estimate.  The idea of a stretched Centaur shouldn’t be too crazy when you realize how many iterations General Dynamics, Martin Marietta, and Lockheed Martin have done on the Centaur just in the past 20 years (including 5m diameter Centaurs for use on Titan IV among other things).  The 1750lb for the extended mission kit is also based on numbers from previous papers LM/ULA has published about converting their stages over for longer-duration missions.  Total dry mass I assumed was 9850lb.  Note that the Atlas V 552 performance numbers also include 5400lb worth of Centaur burnout weight, so you only have to provide ~4450lb worth of “payload” for the Stretched Centaur.  Also note, that if you tank the stretched Centaur up all the way for launch, it should probably increase the payload capacity of the Atlas V 552 a little compared to a normal Centaur, but for purposes of this analysis we’re assuming only the nominal payload of a normal Atlas V 552, to be conservative.
• Single Stage Lunar Lander/Ascender: This stage takes the crew the rest of the way to the lunar surface after the Centaur has provided the first part of the descent burn, and then provides the ascent burn, and the burn to take the crew to the L2 staging point to rendezvous with the Command Module.  I budgetted 1100m/s for its portion of the descent burn, 100m/s to allow for a 90s hover to find the best landing spot, 2650m/s for the lunar surface to L2 burn, and about 50m/s more for contingencies.  This is probably the most aggressive part of the mission.  For this vehicle, I’m assuming a piston-pump-fed LOX/CH4 stage, based off of the piston pump and LOX/Methane engine work XCOR has done  (possibly combined with stuff that we at Masten have done that they haven’t like gimbals, throttling, etc).  The piston pump requires very low net peak suction head, which allows for very low pressure tanks, that can be made of  the LOX/Cryo-compatible Nonburnite composites that XCOR has been devleoping.  XCOR developed the piston pump and Nonburnite composites explicity for making propellant tanks out of shapes that aren’t typical for propellant tanks (in their cases to make the CG numbers work, they wanted to do LOX-filled “wet wings”).  Using this technology, instead of heavy pressure fed tanks and heavy helium tanks, you have lightweight composite tanks that can actually form part of the load-bearing structure of the vehicle.  As I understand it, based on my recollection of their public statements, the piston pumps they’re looking at using scale to about enough flow for a 2500lbf engine in a single pump.  By combining them with the 7500lbf engine XCOR developed (with a nozzle extension of course), you have significantly more thrust than you need for landing.  More importantly, you can possibly make the three pumps operate in a redundant fashion, so the loss of one pump can be tolerated at any point in the mission, and the loss of a second pump can be tolerated through most of the mission.  If done right, the pumps could be “armored” as XCOR calls it, but placed in such a way that they have removable manways between them and the main compartment that would allow for shirtsleeve troubleshooting/repair (the pump compartments would need to be done in a manner that if something went horribly wrong, that any debris/blast would be directed away from the crew cabin…but I can imagine a few ways that could be done).  All told, I’m assuming a 4350lb dry weight, a 9000lb propellant weight, 500lb worth of hardware to be left on the moon, and a 360s Isp.  The LM ascent stage was 4200lb, but held only 65% of the propellant mass, and only about half the propellant volume of this lander, and didn’t have to do landings, and didn’t have to support the crew for as long (about 3 days vs. the target 9 days to give you a week on the surface and 2 days in transity to L2).  But as mentioned above, it used pressure fed tanks, with the mass of a helium blowdown system, had to provide significant RCS capabilities since the stage did not have a gimballed main engine, was using crappy 60s era electronics and electrical systems, and had tanks that were entirely non structural, and also didn’t have access to modern materials like lithium-aluminum or modern composites.  However, the 13,850lb total mass for the lander actually compares pretty well with the 13,510lb currently assumed for the pressure-fed, hypergol-fueld Altair Ascent stage (from this document), which carries 4 crew for the same mission duration.
• Pre-Depot LOX Tank: This ~2.2klb Tank holds ~57.1klb of LOX for the Stretched Centaur.  It includes a docking port (possibly using LIDS technology?), a sunshield, and a Centuar-derived LOX tank.  It gets launched as the sole payload for the Delta-IVH, using up all but about 200lb of its capacity.  But since it is so dense, it might be able to get away with using a shorter (and lighter weight) fairing than is typical for Delta-IVH if that wouldn’t require lots of expensive aero analysis.  This tank, if launched with the LOX pre-chilled can hang out for over a month waiting for the Atlas V 552 launch.
• LH2 Drop Tank: This ~62.5 m^3 tank weighs about 2000lb (with another 2000lb budgetted for connecting structures between the various parts of the launch stack).  It would be housed between the Lander and the Command Module on the Atlas V 552 launch.  It would possibly use 5m tankage derived from the Delta-IV US.  After reaching orbit, the LH2 from this tank would be transfered (using propulsive settling) into the Stretched Centaur.  After the Command Module docks with the Pre-Depot LOX tank, and has transferred all the propellants from that (and discarded the pre-depot LOX tank), the CM and empty LH2 drop tank would separate from the stack, the drop tank would be discarded, and the CM would reattach to the lander much like was done on the Apollo Missions.

Now, this mission model isn’t perfect.  It uses most of the capabilities of the two launchers without a huge amount of margin (except in the fact that the Atlas V 552 with stretched Centaur probably has some margin built in that isn’t being explicitly called out).  And I’m not a fan of launching the crew on an EELV with 5 solid strapons.  It would be a lot easier if you assumed the development of something like the Common Upper Stage that ULA has been talking about recently.  With that, you would have tons more margin (since a CUS would add nearly 7mT of capacity to the DIVH, and probably at least 5mT to the Atlas V 552–possibly enough to go with less or no strapons on the crew launcher).  But it demonstrates that a 2-launch EELV mission using almost no modifications to existing launch vehicles (beyond the Centaur mods) is within feasibility.

The system also has several good things going for it.  First off, it can deliver lunar crew to the surface without a depot.  It doesn’t need Autonomous Rendezvous and Docking (since the rendezvous and docking can be piloted), or tankers to be developed.  It doesn’t need HLVs or 10m fairings (everything can fit within a stock Atlas V fairing).  It doesn’t need really long term LH2 storage in orbit.  It only requires two launches for the mission, and doesn’t put anywhere near as much launch timing constraints as the ESAS architecture does.  It can provide for cargo missions (~19klb delivered mass to the surface assuming that 2klb of the lander stage is in the form of a removable crew cabin, which just happens to be enough to land a Bigelow Module).

And most importantly, if depots do come into existence, it can immediately take advantage of them.  With just an LEO depot, you can both cut down on the number of EELV launches to just one (and use lower-cost systems like Falcon 9’s, Zenits, Ariane-Vs, Soyuzes, future commercial RLVs, etc to launch the remaining propellant).  Also by getting rid of the huge LH2 drop tank, you simplify the stack, remove about 15klb worth of hardware from the Atlas stack , dropping it to the point where it can possibly be launched by a 502 launch instead of a 552 launch (since the stretched Centaur provides almost as much propellant as a Phase 1 Atlas, which was supposed to boost the LEO capacity of the single-stick Atlas to almost 30klb).  Or you could use that saved mass to beef up the lander and/or command module for more capable missions.

If you have both a LEO and an L1 or L2 depot, the Centaur can top itself up again that depot, and provide a much larger chunk of the descent burn to the lander stack.  With enough propellant left over to return to LLO then to L1/L2 after separating from the lander, allowing the Stretched Centaur to be reused multiple times.  With such a system you could actually soft-land bigger payloads than the Altair cargo lander…and you’d have the capability of making the lander and transfer stage fully reusable.  The transfer stage, since it wouldn’t see atmospheric flight, reentry, lunar dust, or even particularly bad thermal environments should actually be reusable for several flights–the RL10 is after all rated for 200 relights.  The lander may be tougher, but by the time you have an L1/L2 depot, you’ve probably had enough time (and enough surface infrastructure built up) that you can work that out to.

Ok, so maybe it’s not so bad of an idea after all.