Selenian Boondocks

This one is in response to an ongoing discussion about the pressurized volume in the various deep space exploration architectures. Apparently an Orion will provide  sufficient crew volume for extended missions while a Dragon will not in many peoples minds. The suggested often remedy is a Bigalow inflatable module for the extended missions.

Henry Cate has a cheaper and faster solution. The trans-Lunar/Mars/asteroid injection stage has a large hydrogen tank. Add a small header tank to the vehicle to store the residual hydrogen in the main tank and vent that main tank to vacuum. Fill the tank back up with breathable air and dock the Dragon to it as in Apollo/LEM. Requires compatible hatches designed into the tank.

The living volume gained will necessarily be Spartan for the same reasons that Wet Skylab was rejected. It does however provide at least 14 cubic meters of volume for each ton of hydrogen that was carried in the original tank. This in addition to the usable Dragon volume would quickly dwarf that of Orion without the launch or purchase expense  of the inflatable module.

With the upcoming attempts to simulate a landing over water with the Falcon IX first stage comes a possible opportunity to expedite actual re-usability.  It just might be possible to grab a stage that has reached zero velocity over the water with a helicopter and return it to a barge or dry land for inspection and possible early reuse.

My understanding is that over the next several Falcon IX flights SpaceX is going to try to reenter the first stage and bring it to at least a momentary hover above the Atlantic before the stage goes swimming. Saltwater immersion is unlikely to result in a re-flyable stage even if it enters the water at near zero velocity. The recently fired hot nozzle will probably object among other things less dramatic.

Landing on a barge has been suggested several times and is even the target of a patent attempt by Blue Origin. In this early case, a barge may not be the best choice. The barge would either have to be exactly where the stage reaches the surface or the stage would need enough propellant to ensure reaching a possibly out of position vessel.  The landing gear would have to be fully developed for the first landing (barging?) attempt, which is a mass and complexity issue. The tall and skinny visuals of the Grasshopper would concern me for a barge landing in anything other than a perfect sea state.

Starting with attempt one at reaching near zero velocity water relative with a Falcon IX, it might just be possible to rendezvous with a helicopter before the stage goes swimming. The factor of ten or so speed difference between a chopper and a barge might make the difference between possible and not for the vehicle meeting at the precise time of hover. The first rendezvous attempt might be no more than a tag-you’re-it with a fast maneuverable bird with a light line with fail-safe breaking strength. A momentary hook up with controlled line tension before the connection is released or breaks. It may not be possible to perform the rendezvous in this manner which would make a full blown recovery attempt questionable until a few answers are obtained.

Even before attempting a full stage rendezvous, it would make sense to try this with either the Grasshopper or a vehicle supplied by others like Armadillo or Masten. An over land demonstration with known vehicles would seem to be a low risk method of training pilots, rocket controls, and developing effective rendezvous techniques. The first trip several hundred miles off shore with few witnesses should not be the first try at a recovery.

If all the preliminaries point to a useful possibility of recovery, then it might make sense to go get a real one. I have no idea of the possible attach points on a Falcon IX first stage in simulated landing configuration. There might be something in the inter stage area that would be just fine for a simple hook from the chopper to snag, or there may not. If there is not, it might make sense to hang a cargo net under the chopper and simply net the stage. It seems possible that loads could be distributed in such a way as to eliminate damaging stresses on the stage so that it could be reused without rebuilding.

The hooked or netted stage could be gently lowered into a soft cradle specifically designed for the stage on a barge or land within range of the helicopter. With air refueling it might be possible to return to launch site for fast turnaround after some experience is gained.

It might just be possible to reuse the Falcon IX first stage earlier than commonly expected without even the mass and cost of landing gear. This would certainly be performance enhancing compared to a full return to launch site vehicle requirement.

These ideas are not all mine or all new. I just thought that it was a good time to mention them again.

I made the mistake of reading about a Senate Commerce “hearing” held today regarding NASA’s human spaceflight plans. While some of the points made I actually agree with, one of the witnesses (Steve Cook of Dynetics) made an argument that I think merits some skepticism. The argument, which you’ve likely seen a lot recently, goes like this “If only NASA had a stable long-term exploration program, with established destinations and dates, the private sector would be jumping all over itself to create business plans supporting NASA’s exploration efforts.” At its core, this argument and others like it seem to imply that if only we hadn’t cancelled the Constellation Program, everything would’ve been better.

Of course, maybe I’m wrong. After all, don’t you remember all of those startups who were champing at the bit to support NASA lunar bases, which were totally devastated when Obama cancelled Constellation so he could funnel money to his corrupt, moneybags campaign contributors? I can still hear their cries of frustration at how all the billions of dollars in VC money they had lined up to march in step with NASA’s glorious Apollo rehash was now going to be wasted.

[Lest any readers who don't know me better think that I've decided to take advantage of recent Colorado referendums regarding recreational herbage--that was bitter and disgusted sarcasm.]

More seriously, I’m curious where all the “Consistencyists” were when Mike Griffin came and completely changed course on how to implement President Bush’s VSE. I may have missed it, but unless my memory faults me, most of the loudest proponents of the importance of consistent long-term plans were perfectly fine with Griffin gutting the HR&T technology development programs, microgravity science funding and ignoring all the work that had been done by industry to-date on developing a lunar return project that leveraged existing launch assets and focused on actually going to the Moon instead of trying to force NASA to stay in the launch to LEO business.

If Congress really wants commercial industry to more actively engage themselves in NASA’s exploration efforts, and to invest private money in ways that are synergistic with their exploration goals, they could try establishing a realistic long-term plan for NASA that fits within realistic budgets, they could try to seriously work with industry to understand what markets industry sees as viable that are also synergistic with NASA’s exploration desires, and work with them to retire the technological risks that are impeding the commercial development of those markets. Leveraging existing, underutilized commercial launch systems and space vehicles such as EELVs, Falcon, Antares, Dragon, Cygnus, etc for placing exploration elements, propellant, and crew in orbit, instead of wasting most of NASA’s exploration budget building unnecessarily duplicative systems to launch stuff into LEO might also help. Ignore those elements of budgetary realism, actually working with industry, and stopping the cycle of blowing all your money on LEO launch systems when commercial alternatives exist, and all of the destinations and timetables in the world aren’t going to attract honest commercial interest and support.

To paraphrase one of my favorite posters from Despair.com, consistency can actually be a virtue, but only if your aim is actually on-target.

I’ve been following the various twists and turns in NASA and private interest in NEOs for a few years, ever since the Flexible Path concept was suggested by the Augustine Committee. While I’ve always been of the opinion that the Moon makes the most sense of all the potential off-earth bodies to settle and develop, I at least could see the appeal of the original Flexible Path concept. Basically you stagger the development of systems in a way to keep the peak “cashflow” needs for NASA as low as possible, and try and do some useful missions as soon as enough mission elements come online to support those missions. At least at the time, that seemed to imply that going to NEOs before the Moon looked feasible.

But somewhere along the process, that idea seemed to go off the rails. Instead of NEOs being a quick “target of opportunity” that could be visited cheaply along the way to the Moon and eventually Phobos, Deimos, and Mars, you started seeing concept architectures coming out of NASA for these massive NEO mission stacks complete with four or five new pieces of expensive in-space hardware that needed to be developed (a Hab module, an MMSEV, a CPS, a big solar electric tug or two, etc, etc) just to visit a NEO. At first I was somewhat baffled at why NASA thought you needed a significant portion of a Mars mission stack to visit a NEO. After all, Josh Hopkins of Lockheed Martin had suggested a much more modest approach (Plymouth Rock) using just two “kissing Orions” that seemed to close just fine.

After doing a little digging, and talking with Josh and others involved in looking at NEO missions from outside of NASA, the best answer I could get was that NASA had set some rather arbitrary minimum requirements that were driving their mission analyses. One of the most important of these was the minimum size of interest for NEOs. Simply put, they didn’t want to go to too small of a NEO, for fear it either wouldn’t be scientifically interesting (yeah right), or because they were somehow afraid that people wouldn’t be impressed to see an Orion visit a rock not much bigger than itself (more likely reason IMO). I can’t remember what the minimum size they set was, but I think it was somewhere in the 60-100m diameter range. The problem is that for NEOs, that narrows things down a lot, which makes mission opportunities much less frequent, and also eliminates many of the easiest to visit NEOs. Hence the big mission stacks. Not only that, but the larger the NEO targets you look at, the fewer “undiscovered” NEOs there are out there to potentially increase your number of potential targets–we know where most of the biggest stuff is, it’s the sub 100m ones that we only know a small fraction of.

Anyhow, we’re three longwinded paragraphs into this blog post, and some of you are probably wondering “what’s Jon’s point?” The point of this blog post is merely to point out the amusing fact that while before NASA wouldn’t even consider sending a mission to a 20m diameter NEO, the new plan would have them visiting a 5-7m diameter rock instead. While I’m not at all opposed to the new “fetch an asteroid” mission (other than the fact that once you’ve brought it back to L2, justifying using Orion/SLS to visit it seems wasteful), I just find it amusing that all of the sudden 7m is exciting, edgy, and NASA-hard, while before it was well below the threshold of interest. Not even yawn-worthy, really.

I guess I’m just curious what NASA could’ve done if they hadn’t originally restricted smaller NEOs from consideration when doing NEO visit architectures. I’m glad they’re finally reconsidering that arbitrary assumption though.

That’s all.

Powering a rover through the two week lunar night is receiving a bit of attention. Most of the attention seems to be on better batteries and thermal control with some suggesting that nuclear is the only reasonable option for such a long time period of extreme cold and darkness. It may be possible to power a lunar installation through the night without using extreme storage or the dreaded nuclear. It may be time to think about the advantages of a Lunar SPS to bridge the gap between the periods of abundant sunlight energy on the ground.

While Lunar stationary would be restricted to orbits so far out as to make the required rectennas too large for rovers and other small early prospecting missions, a lower Lunar orbiting SPS would be able to service much smaller installations. Required rectenna size is inversely proportionate to distance from the power source. Half the distance being half the required diameter and a quarter of the area required. A rectenna on Earth receiving from GEO needs to be a mile or more in diameter with the beaming distance of 22,300 miles or more if not directly under the SPS . A Lunar SPS at 250 miles or so altitude could efficiently hit a rectenna 1% of that or around 50 feet in diameter. Fifty feet diameter is possibly  reasonable for a small rover on a world with  light gravity and no wind. It is certainly reasonable for a stationary prospector of small size.

The drawbacks are both the intermittent nature of the available power and the variable angle with distance of the rover from the SPS. The intermittent nature of the available power will probably be solved for powerless periods of two to three hours followed by intense charging periods with relatively small battery systems.  Rectennas can be designed to accept power from multiple angles with relatively minor design work. The variable distance means that much of the low angle power will be wasted unless a much larger rectenna is built into the surface station. That waste could easily be more acceptable than nuclear depending on the political environment during a particular mission or exploration era. Even wasting 90% of power could be more acceptable than ground systems that must shut down for half of their working lives.

Once a Lunar SPS is operating in this manner it can be used for a multiple of ground stations accross a period of decades.

I’m flying into Phoenix at 9:00 AM on Thursday 11 April for Space Access. If anyone is leaving the airport for the conference about that time I would be glad to pay for the ride.

hare.john@rocketmail.com

863-206-0794

I finally did a kludgy  test Saturday with one of my engine concepts. Several years ago I posted some ideas here about pump impellers that were also regenerative cooled turbines running a full flow staged combustion cycle. I kinda sorta did that. I used a fire extinguisher bottle combustion chamber with nitrous from a 15 pound bottle from the local speed shop. I ran a gas oil mix (demolition saw mix) through the impeller/turbine direct into the precombustion chamber so that the burn would all take place before spinning the turbine/impeller. What would be the main chamber in a normal rocket was just large enough to neck down from the two inch  diameter turbine discharge to the half inch throat. L* about four as best I can figure. The nozzle was a two/notch/twenty cone as I couldn’t get a bell to shape right on the lathe.

Test set up would be funny to those of you doing real testing. A four foot long plank hinged on one end and hanging from a hundred pound fish scale on the other with the engine up side down in the middle. Any thrust would be double the fish scale reading. The fire extinguisher preburner hung below the board with the nozzle above and facing up. The nitrous fed in through the bottom of the preburner that would be the top of a functioning rocket in proper orientation. The oil mix in the gas was to lubricate the turbine/impeller bearing. The gas because I couldn’t get diesel to ignite the weekend before. Ignition was an Estes C motor. The only cooling was the film from the fuel on the preburner walls. The chamber above the throat and the nozzle was fiberglass over regular steel for ablative cooling.

The nitrous read 825 psi on the gauge that came  from a mechanic friend of mine. In a perfect world I could get about 750 psi in the preburner and 500 or so in the regular chamber at the start before pressure dropped except that I didn’t have gauges in either place. The gas mix was at 10 psi to start with compressed air in that tank. This was all in a pit dug with the excavator out in the woods. The property owner thinks it’s all a big joke, probably right.

Henry turned the gas mix petcock with a long stick through the dirt bank just after I hit the button for the Estes rocket and the switch for the nitrous valve. There was a loud pop and a kind of weird roar and whistle that must have been the turbine spinning up. My guess is about a second before a steady roar too loud for the earplugs at fifteen feet and at least six mach diamonds going about fifteen feet straight up for a few seconds before the burn just stopped. There was still a scream that must have been the turbine driven by the nitrous that was still flowing before I realized I had the nitrous switch. I don’t think there was any combustion instability at all.

James read 70 pounds on the fish scale during the early burn which should mean 140 pounds of thrust. If that is right, then we had something like 600 psi at the throat after driving the turbine. The whole contraption weighs just under five pounds so the T/W would be just over twenty eight which would beat the V2 engine if this were flight ready. The pump must have produced at least as much pressure as the nitrous with a weight of just three ounces for the rotating parts and four more for the housing. The nozzle seemed to keep the flow attached to the far wall for the whole burn.

The fiberglass didn’t protect the steel in the chamber and nozzle with a piece of the steel looking like a cutting torch had hit it for a second or so. The turbine/impeller steel had minor blacking on it that rubbed off on the finger. The extinguisher preburner looks strange like it got hot near the nitrous port but not in the rest of it. The burn apparently stopped when the pressure in the gas tank got too low to keep the pump primed, we are just lucky that nitrous didn’t get into the gas tank. Note to self, check valve might appease Murphy as next time we might not be so lucky. The pump was the risky part and it was the only thing that worked exactly like it was supposed to.

None of us know how to rig data collection to a test rig. We also didn’t set up a camera as we didn’t really expect it to work as it did, so I don’t have any pictures or hard data to post. We are sure enough of the results that we are ready now to pay somebody to do that for us.

The turbopump thingy can easily incorporate a nitrous pump, so it seems possible that we can have a full flow pumped system by this summer. We need to build a regenerative lower section as we seem to be running too much heat flux there. The upper (lower in this test) preburner needs a much smaller volume so that there are no hot oxidizer sections without the fuel film cooling like in the extinguisher bottle.

We should be able to get the weight down under two pounds with material selection more for performance than cheap or free off our shop floor. That would give us a T/W of 70 pumping both propellants. The pump seems like it should be able to get at least double the pressure that we got on this test, which would give us a T/W of 140 in a pumped engine in a perfect world. Fingers are crossed. The whole layout would be much easier if we could afford to build it bigger.

Yesterday I spent Easter with the grand kids. I tried bragging about the rocket but they wanted to see it fly and couldn’t get interested in something that just sat there and I wouldn’t have let them get close to the test anyway. My son had what I would call tolerant amusement at my “hobby”. I’ll bring pictures of the rig and parts to Space Access as this time I have something to show.

Launching material off the moon for use in space is a problem in itself. Many methods are feasible technically. Far less are economically sound. Sending the propellant from Earth for instance is something that should only be done for the very early missions for extreme value payloads like an astronaut or a pure helium 3. The more that can be done with an in situ technique the better. Especially if it can be done on the real cheap, relatively of course.

Launch cannons are discussed for Earth launch on occasion and dismissed just about as often. Lunar gun launch is a different story. With an extruded tube twenty kilometers long it is possible to reach Lunar orbital velocity with accelerations averaging under ten gee. One bar of pressure could accelerate three tons at ten gee with a tube two meters in diameter. More or less pressure  would be used for different masses with a thirty ton payload only needing ten bars pressure to hit the ten gee average acceleration.

Two hundred meters of launch tube would be sufficient for hardy payloads capable of tolerating thousand gee accelerations. Early in colony development chunks of frozen oxygen might fit that bill with the metal containers as part of the product delivered to a Lunar orbiting factory. Firing the gun against vacuum instead of an atmosphere really allows guns to show their best side.

All that is hardly new. One thing I haven’t seen as much of is methods of conserving the drive gas from the launch. I suggest that a reusable sabot be part of the standard equipment to support the payload and seal against gas leakage. That also has been thought of. I don’t know of any proposals to conserve all the drive gas for use again and again.

What may be original is the concept of using electromagnetic braking to stop the sabot inside the tube after it has imparted sufficient velocity to the payload. Electromagnetic braking can be done without the massive electrical generation equipment of the various mass drivers often suggested. It can in fact generate electricity while stopping the sabot. The massive power generated during that braking second or two should be usable somewhere in a colony that is still under construction.

The drive gas will start to be evacuated from the gun barrel into storage tanks as soon as the payload reaches design velocity. The gas scavenging  should continue as rapidly as possible to relieve pressure on the sabot that hopefully extended the mechanical brakes as soon as possible to reach a complete stop well before the end of the gun barrel. The scavenging tanks should have large volume at low pressure to evacuate the barrel quickly. The drive tanks should be high pressure and low volume to facilitate good acceleration control. The sabot is returned to the gun breach area after the barrel has been evacuated to set up for the next shot. The drive tanks can be repressurized from the  scavenge tanks at leisure as long as it is quick enough for the next shot.

The extrusion process should be able to turn out a gun barrel of continuously increasing length early in a colony development. As soon as the first two hundred meter section can handle ten atmospheres it could be used to send 300 kg payloads to lunar orbital velocity. The payload would either need to be caught into a circular orbit or provide a circularizing burn in order to avoid lithobraking at the launch site. The payloads would gradually increase in mass as the barrel extended until reaching the 20 kilometer length that would allow the sabot four kilometers of braking room after loosing the payload.

Muzzle loading air rifle for lunar launch with 100% volatile recovery. Low tech and cheap enough enough for you?

There is much discussion on the best habitat to deliver to the Moon. Dragons, inflatables, and modified propellant tanks as well as methods for renovating lava tubes are really just a start on the discussion. There have been a number of thoughts on some form of in situ construction as well primarily of the tunnel digging variety. This is an in situ materials building concept.

Start with an exterior mobile form for the outside of a tube similar in thought to a concrete slip-form. It is sketched in blue in the cartoon. The red section at the bottom of the structure is the heat source that vaporizes the lunar regolith so that the vaporized iron rises to condense on the inside of the slip-form in a vacuum deposition building process. The liberated volatile gasses build up in the completed tube until enough concentration exists to compress them into marketable  volatiles in the vicinity of the airlock pictured in green. With the pressure in the extrusion section never exceeding a fraction of a psi, the slip-form will be able to move on when the shell is just millimeters thick with the remainder of the vacuum deposition to be done against the incomplete tube.

The airlocks are moved toward the construction zone as the tube extends onward. Every time the airlocks move forward there is more living and working space for the Lunar settlers. Only after the metal is thick enough to handle the weight is a few meters of regolith piled on it for radiation protection and thermal control.  It seems just possible that the vacuum deposition tube extruder could be debugged on Earth and made simple enough to require very little human maintenance and supervision. It just might be possible to prebuild living volume robotic-ally before the first settler even lands if the concept could be made simple enough.

Three dimensional printing has been getting a lot of attention lately for building on the moon. I think the printing will be more effective for furnishing the finished tube quarters leaving the heavy construction to simpler and cheaper brute force methods. Once one tube extruder has been shown to work in real lunar conditions, the printers might be best used producing more extrusion slip-forms and power sources in a Von Neuman scenario. Lunar colonists could be faced with an embarrassment of volume riches if the machines are effective enough.

Extrusion of planned settlements could follow whatever plan seems best for the uses intended. One long tube might be considered best in one place while a snail spiral might be preferred somewhere else. One group might best be served by a main street central tube with others laddering off at intervals while another might just think in terms of a wagon wheel layout with all traffic passing through a central hub. tube roads might be feasible if sufficient speed of construction is demonstrated.

The heat source to vaporize the regolith inside the slip-form area can be nuclear or solar depending on the available technology and the political climate when the project is started. If solar, construction would be during the Lunar day by necessity. If nuclear, then night construction might be best so as to be able to radiate the excess heat that would be a necessary byproduct of this construction technique.

Solar powered construction in the daytime would have an enormous waste heat problem without normal convection or fluid conduction available to carry off the excess. It might be possible to shade local areas of regolith from the daylight in order to preserve the cold temperatures reached during the Lunar night. These protected cold traps could be used as heat sinks for the excess heat transported from the building projects. Better would be if the excess heat drawn off from the construction could be converted back into electricity that could in turn vaporize more regolith for the extrusion. Even a 10% efficiency would reduce the required cold trap volume and the required solar collection requirements.

The volatiles liberated during the construction would be possibly more valuable than the structures themselves for both rocket propellant and living requirements. The market for volatiles may even drive the trades toward super heavy tube walls that are just a byproduct of the industrial extraction of valuable export material.

If this were to work on the moon, it could obviously be adapted to Mars and the asteroids to create in situ buildings throughout the solar system.

Just a quick thought for fun.

Suppose some execs got together and had a solid commercial business case for heavy launch ASAP. ATK supplies well tested Shuttle four segment SRBs to mate with current Falcon 9 equipment. Two SRBs are mated with a Falcon 9 core and two strap on stretched  tanks built with Falcon 9 tooling.

At launch the assembly weighs more than the Shuttle stack did and has lower acceleration than the traditional stack even with the Falcon 9 core thrust. At SRB burn out, the falcon 9 and two partly empty strap on tanks mass about 2.4 million pounds and are on a lofted trajectory to compensate for the less than 1 T/W ratio. From the 1,500 m/s or so velocity at separation, the mass ratio to orbit should be around 8. Considering the strap on tanks are dropped three or four minutes after SRB separation, mass in orbit should be on the order of 280,000 pounds.

Your 140 tons is in orbit. How much of it would be usable payload?