Selenian Boondocks

guest blogger john hare

What if games can be quite entertaining even if not practical. This particular one is what if Griffen had dictated an RS-68 for the Ares? It is existing and has considerably more thrust than the J2S, which would seem to imply a more capable second stage with considerably more payload to orbit.

Second glance is where the problems and fun start. A fifth segment was already required for the Ares I first stage even with the available thrust of the J2S, so the SRB would seem to be even more inadequate to support an RS-68 upper stage. Unless you parallel stage to get enough take off thrust, but then you are stuck with a clumsy layout and a sea level  nozzle on the upper stage. Serial stage seemed to be a requirement. Using a pair of stock SRBs would provide enough performance to lift a large upper stage compatible with an RS-68 fitted with a vacuum optimum expansion nozzle, but that would have been a different game altogether.

This what if idea comes from another direction. What if the gas generator cycle RS-68 pumped it’s propellants into the SRB to increase it’s thrust as much as adding another segment only without adding the mass and development of that segment? The gas generator cycle presumably can send propellants through a pipe without concern as to where they are actually used. So the plumbing for the upper stage has two flow paths for the propellant down stream of the pumps. One goes to the first stage  SRB to boost thrust and ISP, while the other path goes to the RS-68 thrust chamber as second stage propulsion.

With the hundred foot L* of the SRB and the rough and tumble combustion of the solid, it would seem that there would be no problem with mixing and burning even with minimal injector capability. A dozen or so ports of inches in diameter should be sufficient. The effective sea level Isp of the virtual RS-68 should even be higher than a stock version as found on the Delta IV because the expansion ratio would be less and the exhaust temperatures higher due to the much higher temperatures of the solid combustion products. The H2/O2 combustion would actually lower the temperature of the solid rocket exhaust though which would drop that effective Isp some. The net Isp effect would seem to be similar to a stock SRB parallel staged with a stock RS-68. The total thrust and Isp would seem to be a bit higher than the five segment SRB while being much lighter.

Testing could be by bolting an H2/O2 propellant supply to a stock SRB at ATK’s static test stand. It shouldn’t be more expensive or time consuming that the five segment development. It would also test a possible command throttle capability.

If this could be made to work, it would put a ~700,000 pound upper stage at roughly the same altitude and velocity as the Shuttle stack at SRB burnout. A very high expansion ratio RS-68 should get an Isp considerably higher than a stock engine, possibly approaching RL-10 performance. Various assumptions give a mass ratio of 4 to 5 for the rest of the way to orbit. If this different Ares I placed 140,000 to 175,000 pounds in LEO, then effective payload should be in the 30-40 ton class even with the extra tankage supporting the first stage burn.

Properly handled, this would seem to be a better, faster, cheaper way to get a strong medium lift. We all know better, which is why this is just a what if post for fun. Unless the concept itself is viable and can be applied to other vehicles later on.

I told the story of how I had gotten involved with the JSC study of an artificial-gravity/nuclear-electric propulsion (AG-NEP) Mars vehicle study. I came into the study near the end (January 2003) and right before the Columbia disaster.

As near as I could tell, after Columbia happened, nobody kept working on the AG-NEP design, or even on Mars studies for that matter. If they did, I certainly didn’t know about it.

But for some reason, the whole idea kept rattling around in the back of my head. There were a few reasons that the JSC guys had given me that were compelling for AG-NEP as a Mars vehicle.

1. You solve the muscle and bone loss problem through artificial gravity. You don’t have to worry about hours of exercise or fret whether their bones will snap when they re-enter the Earth’s atmosphere. They’re going to be good and strong when they get home because you made sure that their bodies felt a normal level of gravity throughout the trip.

1a. Because you’ve solved the muscle and bone loss problem, the pressing need to fly the mission quickly is tremendously diminished. You can go to Mars and come back in the three-year time frame that is more astrodynamically “natural”, in other words, the time frame that aligns with the Earth and Mars’s movements around the Sun.

2. By using nuclear-electric propulsion, you actually have a credible propulsion system to execute a mission abort if you need to, for some reason along the way. You’re not going to get back quickly, but you can get back.

3. By using nuclear-electric propulsion, you actually have a credible story for vehicle reuse. You could refuel the vehicle and go again. Or you could go somewhere else like an asteroid. You have a lot more flexibility than in a chemical or NTR vehicle.

I liked the basic idea. Here was a vehicle that might actually be a true “spacecraft” as we like to think of them, with the ability to go and come from a variety of destinations and be reused. I imagined that this might be the kind of vehicle that would be in Captain Picard’s ready-room a few centuries from now as a little model, with him saying, “This is a model of the vehicle man used to explore the solar system in the early days.”

But there were definitely residual technical problems with the design as it stood when I was exposed to it. The biggest one had to do with getting the body-mounted electric thrusters to point in the right direction as the vehicle moved around the Sun, and the problem got so bad when you got to a spiral-in, spiral-out scenario around a planet that it was practically a no-go. It came down to the architectural decision to orient the thrusters so that they were firing in the same direction as the vehicle’s angular momentum vector (orthogonal to the rotation plane). That approach certainly solved any problem of plume impingement, but since the inertial direction of the thrust vector was going to change by >180 degrees during the transit to Mars, and by that much or more on the way back, you had to continuously move the angular momentum vector of your spacecraft around, and there was a non-trivial cost associated with doing that. During spiral-in or spiral-out the cost became prohibitive.

The other problem concerned spin up and spin down of the system. The JSC design assumed that spinup and spindown would be done by dedicated thrusters on the habitat module end of the vehicle. That meant a duplication in thrusters and tankage for a capability that you would want to utilize as little as possible.

Despite these problems, I recognized that the JSC design as it stood had also solved a great many problems, and that perhaps it represented a minimum in the design space of overall difficulty. I’m fond of saying “you have to pick your pain” when it comes to system optimization, and that the “best” system always involves residual problems. Perhaps this was as good as it got.

Or maybe it could be even better.

One day I was driving down the street in the pouring rain and a simple sequence of thoughts formed in my brain:

1. I had spent a whole bunch of time trying to figure out how to get solar panels on a MXER tether to point at the Sun while the tether rotated.

2. I had been lucky enough to meet Steve Canfield and had figured out how to use the Canfield joint to fix the problem of pointing the panels at an inertial target (the Sun) while the overall structure (the tether) rotated.

3. The basic problem that the AG-NEP vehicle faced was the need to point its electric thrusters at an inertial target (its thrust vector) while it rotated, much like the MXER tether needed to do with its solar arrays.

4. The reasons that JSC had rejected rotating machinery for the AG-NEP vehicle had to do with the difficulty of moving propellant and electric power across a rotating connection like a rotary joint or slip ring, and these were good and valid reasons.

5. The Canfield joint had no such problems because provided that propellant lines or power cables were flexible, they could transmit fluids and power across a Canfield joint.

thus…maybe a Canfield joint was the answer to the problems of the AG-NEP vehicle!

I couldn’t believe that I had known about the Canfield joint for so long and hadn’t put these utterly compatible ideas together.

If we were to use the Canfield joint on the AG-NEP vehicle, the overall geometry would change substantially. The logical location for the thrusters moved from the center of the vehicle, on a cross-brace, to the reactor end of the vehicle. This kept the high-power lines short since they didn’t have to run all the way to the middle of the vehicle to reach the engines. You could also place the propellant tanks on the reactor end of the vehicle as well.

This in turn led to several other vehicle advantages:

1. The moment-arm from the reactor module to the hab module is shortened (or alternatively the moment arm from the CM to the hab module can be lengthened) because now there is much more mass counterbalancing the hab module. The thrusters and the propellant constitute a lot of mass.

2. The truss between the reactor module and the hab module now doesn’t need any “cross-brace” on it or any other body-mounted structures. It can be a strong but simple extensible structure, like a CoilABLE boom, with nothing more than the power connection between the reactor and the hab module integrated into it.

3. The main thrusters can be used for spin up and spindown operations. By placing them on the end of the moment arm, they now have the ability to change the angular momentum of the vehicle, by simply remaining fixed relative to the vehicle during spinup and spindown. In fact, spin rate can be changed during thrusting simply by changing the fraction of the spin arc during which the thrusters fire.

4. The angular momentum vector of the vehicle doesn’t have to point along the thrust vector (like in the JSC design) but can point orthogonal to the spacecraft’s orbital plane. This means that the angular momentum vector’s direction doesn’t have to be altered during flight. This also means that spiral-in/spiral-out maneuvers at planets are no problem.

5. If you wanted to use the AG-NEP vehicle for asteroid missions, the electric thrusters might even be able to be used as “descent engines” provided some “landing gear” were provided on the habitat module.

6. Propellant could be used for additional reactor shielding during the flight.

Over the years since this realization, I’ve developed the capability to show how such a vehicle might look as it rotates. Here’s the link to a Java code that will show the vehicle rotating along with the Canfield joints. You can click and drag to rotate the view around and zoom in and out with the mouse wheel.

For reference, here’s the original set of slides from JSC describing the problem and their original design solution.

In previous posts I’ve mentioned that when I first got to NASA I worked in the Propulsion Research Center, which was a fun place to work because you got to think about and try just about anything you wanted to so long as you could get funding, and there was this sugar-daddy at NASA named John Cole who would fund all kinds of crazy stuff. I never got any funding from John but my patron was Les Johnson, who was kind of like NASA’s “point-man” on tether technology. After about two years in the PRC, Les told me it was about time to quit fooling around and become a serious-type manager like him, and to come and join him in the newly forming In-Space Propulsion project.

So in the fall of 2002 that’s what I did, and before long I was writing NRAs (National Research Announcements) to solicit universities and corporations to bid on technology work for tethers. We put the first NRA out for tethers and got responses and had a meeting where a committee picked the winners in March 2003. After that things started getting serious. We had real money for the first time to do momentum-exchange tether work, and there were still so many unanswered questions that needed to be solved.

Sometimes fate or luck or serendipity drops things in your lap. In the summer of 2002, I met one of the most clever and hard-working people I’ve ever had the good fortune to meet–Dr. Stephen Canfield of Tennessee Technological University. The next summer he was down at MSFC and I was in the middle of trying to figure out the answer to a very thorny problem: if you have a tether that’s spinning, how do you keep the solar panels pointed at the Sun? My friend Kyle Frame and I would sit in my cubicle for long stretches of time with pieces of paper pretending to be solar panels and pencils and sticks standing in for the tether, trying to figure out some way to do it that wasn’t totally foolish.

One day Steve Canfield stopped in and asked us what we were up to. We described the problem and he asked a simple question:

“Do you care what orientation your solar panel is in so long as it is pointed at the Sun?”

I said no, we didn’t care, and then he showed me something he’d been working on since he was a grad student. It looked like this:

He called it a “Trio-Tristar Carpal Wrist Joint.” I thought that sounded like a real mouthful so I just called it “Canfield’s joint” and eventually everyone (except Canfield) began to call it a Canfield joint. It was kind of a crazy looking thing that you couldn’t figure out what to do with it unless you held it in your hands and started playing with it. Unfortunately, in a blog post I can’t reach out of your screen and hand you your own Canfield joint to play with, because if I could you’d figure out in a few seconds what I’m talking about, but the real magic of the Canfield joint is that you can point the joint anywhere in a hemisphere without winding up anything.

The joint has several parts. There’s the “base plate” which stays attached to whatever the joint is mounted to, like your spacecraft, and then there’s the “distal plate”, which points to whatever it is that you want to point at. There are six legs on the joint, in three units. The joint is called a “parallel structure” because there’s more than one load path for the loads to follow, and this is what gives it its potential strength. Where the legs mount to the plates is a simple revolute joint. I didn’t know what that meant so I asked Canfield and he said that it just meant that it was a simple, one-degree-of-freedom (one way to move) joint or hinge. Where the two legs come together you could have a spheric joint (a ball and socket with two degrees-of-freedom) or you could have three revolute joints in series. That’s what we usually do.

I asked Canfield what the joint was for. He said that he originally wanted to use it to replace the CV joints in cars, since if it had all revolute-joints then it wouldn’t need a boot. If I hadn’t had to replace the boot on the CV joint in my car when I was in college and dirt-poor, I wouldn’t have had any idea what he was talking about, but the loss of money was still burned in my mind, so I appreciated that application.

Well, to make a too-long story shorter, I learned how the Canfield joint worked and figured out how to solve my little problem on the tether. Tell me if you like the result:
Canfield Joint on MXER Tether
Medium View of Canfield Joint
Closeup of Canfield Joint

In several posts now, I have criticized the use of nuclear thermal rocket (NTR) engines. In the case of Earth departure stages, I have shown through mathematical analysis that they either do not have a performance improvement over chemical engines (for the overall system) or that the performance improvement is insufficient to merit the titanic expenditure that would be required to develop them. In the case of a hypothetical Earth-to-orbit application, I have shown that there is simply no hope whatsoever for their use.

My writings have elicited strong responses, both here and over on the NSF forum. People have asserted that I am simply wrong, or that I have gone into the analysis with a bias that has somehow compromised my results, or that I have ignored some obscure advanced version of an NTR that they assert can solve the problems that I have identified. Of these criticisms, it is the criticism that my bias compromises my results that troubles me the most. As an engineer, we want to believe that we are immune to opinions and biases but my experience has been that that is not the case at all. They creep into our judgement, and sometimes they sit flat on our face. I have seen for years now at NASA how even the most clever engineers can be seduced or bullied into accepting terrible vehicle designs, and in a perversion of the Stockholm Syndrome, eventually come to “love” the fatally-flawed design that they would have initially rejected.

I cannot be certain that my biases do not affect my work, but I am always striving to reduce them as much as possible. The most effective way I have found to do this is to get to “the numbers” as quickly as possible. We can talk all day long about how much better this technology is than that technology, but when we get to the numbers we can begin to improve the signal-to-noise ratio of our discussions more quickly than anything else I know.

Let me tell a story about a time when my opinion was changed significantly through solid engineering analysis. When I first got to NASA in 2000, I was part of a study called by a number of different names. Some called it “Decadal Planning”. I thought of it as “go to Mars and back in a year”. Within a few months of getting there, I was running a part of the study looking at propulsion technologies, and we were comparing a number of them. I met a lot of people in the field and began to develop a distaste for nuclear electric propulsion (NEP). The study concluded quietly and in my opinion, was a failure. Some of that was my own fault. But that’s some other post.

Near the end of 2002, my boss asked me to be a part of an MSFC response to a JSC study on a new and different NEP vehicle. As I recall they were interested in launch vehicle options. So in January 2003 several of us went to JSC to talk to them. I was very impressed by what they had done.

Essentially, they had asked “what are the big barriers towards sending people to Mars?” and then “how do we deal with them?” The biggest barrier they had identified was uncertainty about what happens to people after many years in microgravity. So they decided from the outset to design a vehicle that incorporates artificial gravity, thus cutting the “gordian knot” that had driven previous mission planners to “fast” trip times. My own experience on the one-year-round-trip Mars study had convinced me that it was utterly foolhardy to try to go to Mars and back too quickly. But I know the idea still rattles around in the Internet and is carried by the astrodynamically misinformed.

To accomplish the artificial gravity approach, the JSC study anticipated using the natural countermass of the nuclear reactor that would power the NEP vehicle to counterbalance the mass of the inflatable crew habitat. The boom that would be present anyway to keep the reactor away from the crew would now double as the separator needed for artificial gravity. Any time you can get a “two-for-one” value like that in space vehicle design, you want to take it.

The persistent problem in the design was the need to point the engines along an inertial direction while the vehicle was rotating. The JSC planners had rejected the idea of rotating slip or roll rings, for good reason and based on their experience with the ISS. To keep the engines body-mounted and yet pointing along an inertial direction required rotating the vehicle in inertial space, nearly 180 degrees during the transit to Mars. Once they got to Mars, the “spiral-in” proved very difficult, since now they would need to move the rotation axis of the vehicle through 360 degrees on every orbit. They came up with a compromise, called “minor-axis rotation” that mitigated some of the issues associated with this maneuver, but I don’t think they were terribly satisfied with it.

The design wasn’t complete or perfect, but it was a real step forward. And I was very impressed by their willingness to challenge their pre-conceived notions about Mars travel and examine a design completely different from what they had looked at before. And it changed my mind about the value of nuclear electric propulsion technology. In a future post, I’ll describe how a design innovation I came up with could solve the remaining architectural concerns with the AG-NEP vehicle and make it far more feasible.

My contrast, the response of some of the MSFC personnel I was travelling with was not so open minded, at least with regards to launch options. Before we went over to the meeting, we met at a Denny’s for breakfast. There were two senior personnel, a mid-level manager, and me. One of the senior folks (who’s no longer with MSFC) began to lay down the “MSFC position” for the meeting, which he said would be Shuttle-C. I gently began to demur, saying how the use of Shuttle-C would commit us to using the expensive Shuttle infrastructure for decades to come. He quietly but firmly cut me off and said:

“The answer…is Shuttle-C.”

I understood from the tone of his voice that this decision wasn’t technical, it was strategic. Shuttle-C was based on propulsion hardware developed and controlled at MSFC. If the Mars program went forward, if this vehicle was developed, if Shuttle-C was baselined for its launch, then MSFC would be supporting it for many years to come. I looked over at the other senior person and he was nodding his head in agreement. I looked at the mid-level manager and he wasn’t saying anything. He and I were like enlisted men getting our orders from the officers. And so Shuttle-C it was…the only option we were permitted to present to the JSC study team.

As it was, several weeks later the Columbia was destroyed on reentry. Everything changed at the agency and everyone forgot all about Mars studies. But I still remember what I learned and think that there are lessons to draw from it.

A few months ago, I spent some time describing some calculations of payload fraction that I derived to assist in the design of rocket vehicles. My motivation for getting into this type of work came about from my work on the X-33 rocket when I was an intern at the Skunk Works. I wondered how so many people could think that SSTO (single-stage-to-orbit) was a good idea when the mathematics argued against it.

Right after I joined NASA, in early 2000, I was in a group that was looking at some really advanced concepts, and somehow or another, we got looking at using nuclear thermal rockets for an SSTO vehicle. At first blush, the whole idea seems to make sense. Nuclear thermal rockets offer almost twice the specific impulse (Isp) of chemical rockets, and if an SSTO doesn’t have enough Isp with chemical rockets, then surely nuclear rockets must be better, right?

Wrong. Super wrong.

Nuclear-thermal SSTO turns out to be one of the worst ideas anyone has ever come up with, for two simple reasons: hydrogen and the lousy thrust-to-weight ratio of nuclear thermal rockets. Those are the same two reasons that make NTR lousy or marginal for nearly any other space application as well, but this post will focus on the issues surrounding NTR SSTO.

In the case of any earth-to-orbit vehicle, you’ve got to have the thrust to get off the ground in the first place. Let’s assume that we’re dealing with a vertically-launched NTR SSTO. It has to have a vehicle thrust-to-weight ratio greater than one, and probably a fair bit better than that in the first place, just to get off the ground. So we can take those expressions that I derived before, assuming hydrogen as a propellant and the engine thrust-to-weight ratios that have been quoted by NTR proponents like Stan Borowski to quickly try to figure a payload fraction for an NTR SSTO.

We find the propellant-mass-sensitive term (derivation here) assuming the liquid hydrogen has a density of 71 kg/m3, ullage of 3%, a mixture ratio of zero, and a tank structural mass factor of 10 kg/m3. This gives us a value of 0.1452 for this term.

We find the gross-mass-sensitive term (derivation here) by assuming that the engine has a vacuum thrust of 15000 lbf, a weight of 5000 lbm, and vacuum thrust-to-weight of 3 to 1. I’m not even going to “ding” the engine for sea-level performance, since as we’ll see, it won’t even matter. With a vacuum T/W of 3 and the same for the sea-level T/W and an initial vehicle thrust-to-weight ratio of 1.25, and we’ll just say that the thrust structure doesn’t weigh anything either, the gross-mass-sensitive term comes out to be 0.4167.

We’ll also ignore any recovery hardware (wings, landing gear, TPS, etc) and say all that weighs nothing. We’ll assume that the engine has a vacuum Isp of 900 seconds and that it takes 9200 m/s of delta-V to get to orbit. Plugging those numbers in the rocket equation gives us a mass ratio of 2.835 (very good!) and a propellant mass fraction of 64.73%. Next we use the prop-mass-sensitive and gross-mass-sensitive terms, along with the propellant mass fraction to get the payload fraction (derivation here).

We start out with the final mass fraction (1 – prop mass fraction) of 35.27%. It doesn’t get any better than that. Then we subtract the gross-mass-sensitive term (41.67%). Now we could stop right here, because we’re already negative (-0.064). That is to say, even before accounting for the issues with tankage, we’re already out of performance. The engines weigh too much. But we’ll keep going and subtract the product of the prop-mass-sensitive term and the propellant mass fraction (0.6473*0.1452 = 0.0940) and we end up with a payload fraction of -0.1579.

So it’s a no-go with these engines. Our payload fraction is grossly negative and we’ve got nothing. It’s clear from the magnitude of the numbers that the engine thrust-to-weight ratio is the main culprit, although the “fluffy” liquid hydrogen tanks don’t help much either.

So what kind of engine performance would you have to have to get even a zero payload fraction? Well, I ran some rough calculations based on a variety of speculative vacuum T/W ratios for some putative NTR engine, at a few different values of specific impulse and plotted the results here:

The graph tells the story. To get payload fractions of zero (a launch vehicle of infinite size) you have to have a T/W at 900 sec Isp of over 10. That’s more than three times the T/W that Stan Borowski projects for his sporty 15K NTR design, which he says will have a T/W of 3. So if you think that Stan or others can design an NTR that only weighs a third of what he thinks it will weigh, then you can dream about an NTR SSTO of infinite mass.

As for me, I’ve thought for some time that NTR was a really bad idea for almost every application for which it is considered. The SSTO application is probably the worst.

guest blogger john hare

The increasing tempo of VTVL development flights and the recent success of the Falcon 9 lead to possibilities for a different type of cooperative venture. Two companies have VTVLs testing  that are pretty much gas-n-go while SpaceX has vehicles that are quite difficult to get back. Using gas-n-go boosters to improve an expendable rocket payload might be a viable business.

Two VTVLs are built proportionate to the Falcon 9 and used as strap on boosters with propellant cross feed so that the Falcon 9 is fully fueled up when the VTVLs stage off and return to launch site for a pinpoint landing. What I propose that is different than what I have seen before is that the VTVLs separate at about a minute into the flight before transonic flight and maxQ is reached. By separating at high subsonic, the VTVL vehicles and the coop vehicle clusters never have to be designed for or subjected to the stresses of transonic and supersonic flight.

Subsonic flight is a far more forgiving and understood aerodynamic problem than the higher velocities and leads to considerably less problems, though also with less results. By staging at high altitude the Falcon’s engines are attaining near vacuum thrust and the vehicle could be considerably heavier than on a ground take off. A small tank stretch and a subsonic boost should get about 40-50% more payload to orbit. While the Falcon would be the main player, the supporting cast could improve the bottom line considerably with benefits for all.

The benefits for Falcon would be of the same order as the subsonic air launch scenarios that so many have studied. The VTVLs just do the airplane’s job. It is the emergence of the fast turnaround rocket vehicles that make it possible to virtually airlaunch in unlimited sizes.

The VTVL players would have a market for a fairly low velocity vehicle with a high dollar (compared to the suborbital field) market that doesn’t require high flight rates. It would give them early experience with a larger vehicle than would fit in their normal course of development, and a large launch assist platform in the bargain. Though the vehicles developed for Falcon assist would not go supersonic in their booster role, they would have plenty of size margin for modifications to allow them to carry relatively large VTVL upper stages to mach 3-4 and still do an RTLS maneuver for another flight or two that day.

guest blogger john hare

SpaceX nailed the Falcon 9 on the first try. There is enough crow being eaten around the country now that somebody should put out a cookbook. My serving comes from the expressed belief that the opening of space will come through the incremental development with RLVs starting from suborbital through small orbital and so on, and not through all up test flights of ELVs. I could develop a taste for this brand of crow, so SpaceX, serve it up. Prove me wrong.

I haven’t changed my opinion of how space will be opened up in the long term, but this is an increasing sum game. I can cheer for my favorite teams without wishing ill on the others.

Ares is a different team in a different league playing a different decreasing sum game, so I can wish ill on that one. I think their players will be drifting away to more productive sports, and the events of Friday will accelerate that drift.

guest blogger john hare

Recent events make it easier to describe one scenario for getting orbital costs down. This specific example almost certainly won’t happen and just stands in for the dozens of possible ways that orbit could become affordable.

Masten and XCOR have a joint venture for developing a methane lander. Consider the possibility of future cooperation. If in five years Masten and XCOR are operating profitable sub orbitals out of Mojave, it will be past time to start on orbital hardware. If both companies have 300+ kg payload capacities to above 100 km, then the Lynx  will have the capacity to lift one of the earlier Masten vehicles (Xombe 0.5?)to above 100 km. A fairly rapid partially retrograde development could have the small Masten vehicle capable of very long suborbital flights if properly assisted.  Say, Mojave to Kodiak, Oklahoma, or New Mexico. The Lynx lofts the Masten vehicle to mach 3-4 at high altitude and stages. The Masten test article finishes the boost for one of those spaceports while the Lynx returns to it’s runway.

The experience gained in test flights with almost existing equipment will go a long way toward convincing investors that orbit is well withing the reach of a conservative  three stage to orbit transport. (Pete suggested this as two stage over on Transterrestrial) The reentry problems can be explored for mach 10 to 15 including turnaround time and environmental impact. Reentry noise can be one of the largest hurdles for inland recovery of second stages. If, and only if, inland recovery can be done in an acceptable manner will this concept work. A tiny smallsat stage might be launched from the Xombe 0.5 if required to make the point.

For a second set of tests, Masten launches a three barrel vehicle with their full size suborbital craft. The two outers deliver the same assistance to the center vehicle as Lynx does to the smaller craft. The full size suborbital vehicle lands downrange at one of the available spaceports. Once this is explored, a 300 kg upper stage is used to place a useful smallsat in orbit. Some testing is done on recovering the upper stage.

If the above tests confirm that the technical concept is sound and acceptable to the uninvolved public at the recovery spaceport, then funding can be sought for a full orbital system. A much larger XCOR vehicle (Panther  for this post) is developed to launch the full size three barrel Masten (Xorbit for this post) assembly. This time the two outer vehicles cross feed propellant to the center vehicle and land at one of the downrange spaceports. The center vehicle goes on to deliver 300 kg of payload to orbit. This allows XCOR to focus on the much larger suborbital Panther while Masten focuses on improving existing vehicle performance and reentry. The Panther could also be the rent-a-booster as Clark suggests over on Space Transport News. The technical risk to orbit from that point would be comparable to funding XCOR and Masten to a full suborbital operation now.

A launch organization has four main costs. Development costs. New vehicle costs. Fixed operating costs. And marginal costs. Only marginal costs are mostly unaffected by flight rate. On a gas-n-go vehicle propellant is a major marginal cost.

On this three stage to orbit with each stage having a mass ratio of three, (total MR=27) there would be about 36,000 kg of propellant or about $18,000.00 per launch. Other marginal consumables would double this to $36,000.00 per launch or about $120.00 per kg in marginal costs.

Starting from a base of operating vehicles and recent development experience with an intact team, I speculate that the development cost would be about $250M. Expected interest on the money for this high risk field could easily be in the 25% range for a development cost of $65M per year to service the debt. If there is one flight every two weeks, $2.5M per flight to service the debt. At two flights per week, this drops to $625K. With multiple airframes flying twenty payloads per week, this drops to $63K per flight to service the debt.

New vehicle costs for the four airframes (Panther and 3 Xorbits) could be in the $20M range after the bugs are out. At a flight every two weeks, it would take $770K to pay off a set in a year. At two flights per week per airframe, it drops to $193K to pay off in a year. If some financing can be arranged, then costs can drop considerably more. (It could be much less in the reality as the Panther could lift 20+ times per week while the Xorbits could possibly do two each.)

 My first guess on fixed operating costs is $30M per year keeping two or more spaceports fully operating with all support staff, including transporting the Xorbit stages back to Mojave. These numbers are based on this guess. At a flight every two weeks, $1.154M per flight. At two per week, $289K per flight. At twenty flights per week, $58K per flight figuring that staff doubles at that flight rate.

flight frequency   debt    vehicles      fixed              marginal           total      

two weeks           $2.5M    $770K       $1.154M        $36K            $4.46M     

two per week    $625K      $193K         $289K          $36K            $1.143M    

twenty per week$63K      $193K         $58K              $36K              $350K     

If it worked out like this, then a flight rate of once every two weeks would break even at a little under $7K per pound, while twenty flights per week could get it down to around $530 per pound. A lower flight rate under every two weeks would make the plan unworkable while a higher flight rate than twenty per week would make it somewhat cheaper than $500.00 per pound. These numbers assume that interest only is paid on the development and that the vehicles must pay for themselves in one year.

This is very much a first generation space transport. Obvious places to get the cost down further are to reduce development costs while still getting an acceptable vehicle group. Getting a better interest rate. Improving propulsion and dry mass fraction to get the mass ratio down from 27 to 16 or so with less dry mass to lift in the first place. Getting it down to two stages to orbit to eliminate supporting downrange spaceports and the extra stages. Flying each airframe more than the twice a week I have here. Reducing the cost of new airframes and spreading the payback time over more than one year. Reducing the number of support personnel required per flight. And so on to get costs insanely low compared to the present day.

 Developing orbit into an economic driver is going to take a lot of work and intelligent handling. Orbital development to date with existing transportation is a mere shadow of the possibilities given transportation that is economical, reliable, and convenient. Everything changes when a decision can be made and executed reliably and affordable in a matter of weeks or possibly days. When was the last time you scheduled anything years in advance? How many times in your whole life do you schedule something years in advance and know that it might cato before it gets started? That is what the whole launch industry is forced to do at this time.

Basing on this cooperative venture creating a somewhat convenient transportation method to orbit, speculation on markets becomes possible. The first markets are somewhat limited in orbital inclination by the restrictions of the downrange spaceport requirement. The choices are a somewhat retrograde high inclination orbit, or a normal orbit with an inclination at  Mojave’s latitude. Anything to ISS is out due to inclination restrictions. Tourism is out due to the small vehicle size. What can you do?

While proving the vehicles, expensive payloads are out, so the focus must be on things that have low intrinsic value with only being in orbit making them valuable. Propellant is the first obvious choice. The first launch to each available inclination carries the largest propellant tank that it can as payload with subsequent launches filling it up. Or possibly one of the ELV companies leaves an upper stage in the right orbit that can serve as a depot. This is only good if there is a market for propellant from that particular orbit though. I dislike the concept of buying propellant, or water, or sand in orbit for the purpose of artificially creating a market.

The first stopgap depot would need to be in a normal orbit at Mojave’s latitude. When it has sufficient propellant delivered to make it worthwhile, one of the majors launches an un fueled GEO bird from the Cape to rendezvous with it. A 25 ton GEO bird unfueled could take on 50 tons of propellant before boosting to it’s final orbit. This would make for a very massive satellite compared with the current operational GEO sats. This would take about 170 launches of Panther/Xorbit. A contract for a million a launch would be profitable to both sides. The GEO provider would have the capability of a 75 ton launch vehicle for an extra $170M per GEO sat while only risking a 25 ton launcher. The Panther/Xorbit companies would need to get the launch rate above 3 per week to make a profit and pay off their debt. The 3 launches per week to reach profitability would take over a year to complete for each GEO bird that takes advantage of the capability. Assuming there are at least two of these monster satellites per year to refuel,  the launch rate is at nearly 7 per week. If deliveries could reach that flight rate, about 40% of each launch would be for profit or paying off the debt. Debt payoff would be under 3 years and would further reduce the launch costs by a factor of almost two, which would make the launches even more profitable at that price point.

Propellant is not the only possible cheap on the ground expensive in orbit payload though. Military surveillance would benefit from the ability to place fleets of cheap observation satellites in orbit. The gaps in coverage of the existing sats are most likely known to the people with something to hide. They probably time most of their sensitive moves to take place in these gaps. The military could easily spec an LEO surveillance sat that could be mass produced for under half a million. Though these birds would be far less capable than the high end machines they use now, a thousand of them would provide wall to wall coverage. At a million a launch and half that for the actual satellite, they would be looking at $1.5B for the whole program not including collecting and integrating the data. Allowing three years for placing the constellation, flight rates and costs match that of the propellant market only. Under three years to payout and higher profit thereafter.

Communications is frequently brought up as the target for early market. With cheap reliable launch as suggested here, less capable LEO satellites could be used in constellations with spares on orbit and on the ground ready to be launched in days if required. It is quite conceivable that these 300 kg comsats could be mass produced for a half million just as the surveillance satellites are. A billion and a half for a robust LEO comsat constellation of a thousand birds should be a business plan that would close. A three year launch schedule has the same numbers as the propellant or surveillance constellations. Under three year payoff and 40% profit after at a million a launch.

Astronomy and microgravity science and Earth observation satellites should be a market to match the last two.  Though the individual capabilities are less with the smaller satellites and cheaper construction, a cast of thousands could make up the difference. The ability to launch follow up or replacement experiments in a week or less would make commercial use of LEO much more user friendly. Over a hundred countries with thousands of universities and tens of thousands of private companies would be the market this time. With similar launch rates to the other users, the money works out the same.

If all four of these markets were to emerge though, a flight rate of 30 or so per week would drop costs still further. Costs of under $350K per flight would allow a charge of a half million per flight to make over 30% profit available for the debt payoff or investor ROI. It would also create even more incentive for the users of the service to make less expensive satellites for an even lower total cost. $4.5M per week profit would pay off the development debt in just over a year.

The assumption that these satellites would be inferior to the $Dirksen Galactica models launched by the current providers might not be true. The current birds are just too expensive and hard to replace to fail so every possible effort is spent making them reliable. This means literally gold plating some components and using only hardware that has been “space rated”. The time required to space rate components added to the long lead time on the current launch capability means that by the time any component reaches orbit, it is incredibly expensive, and it is obsolete by standards on the ground. The point has been made many times by many people that many of the current satellites, especially comsats,  are so expensive that the cost of launch is not all that important. With replacement launch on demand, satellite reliability becomes somewhat less important. It becomes possible to launch the latest electronic devices without “space rating” them at all. If they survive, they are space rated, if not, they are not. Henry Spenser has noted that the commercial stuff launched in the Canadian satellite that he worked with has lasted for years. How much capability is in $10K worth of May 2010 commercial electronic components components compared to $100M of “space rated” 1995 electronic components?

When failure is not an option, success can get expensive. We have all read that. How many have also noted that when failure is not an option, success also gets much  less capable?

With launches available on demand for half a million, many high risk or complicated capabilities can be tested and possibly implemented. A 300 kg tug could do considerable work in LEO. It is past time to deploy tethers for extensive testing. Debris cleanup vehicles can be tested and then employed. Fairly small service vehicles could do good work in GEO. National Geographic could afford a 300 kg Lunar, Mars, or NEO probe that was launched on one flight and fueled up by others.

It just gets better and the sky is not the limit.

I had to keep this under wraps until this morning, but it’s now formal:

May 25th, 2010, Mojave, CA, USA: XCOR Aerospace and Masten Space Systems, two of the leaders in the New Space sector, have announced a strategic business and technology relationship to pursue jointly the anticipated NASA sponsored unmanned lander projects. These automated lander programs are expected to serve as robotic test beds on Earth, on the lunar surface, Mars, near Earth objects and other interplanetary locales, helping NASA push the boundaries of technology and opening the solar system for future human exploration.

Masten’s award winning automated vertical take off, vertical landing (VTVL) flight vehicles combined with XCOR’s strong experience in liquid oxygen (LOX) / methane powered propulsion systems and nonflammable cryogenically compatible composite tanks, brings to NASA a powerful and competitive combination of innovative talent with a proven record of producing exceptional results quickly and affordably.

Last October, Masten won the $1 million first prize for Level II of NASA’s Lunar Lander Challenge, beating out a host of New Space rivals, and demonstrating they are the leading VTVL development group in the country. In 2007 XCOR Aerospace’s LOX/methane engine, developed for NASA, was named by Time Magazine as one of the “Inventions of the Year”, recognizing XCOR’s successive advancement in the state of the art of both pump and pressure fed reusable, throttle-able rocket propulsion systems. XCOR and Masten have also demonstrated the ability to rapidly take from concept to live fire, new propulsion and control system designs using innovative rapid prototyping techniques that surpass client requirements in much shorter periods of time than traditional aerospace methods.

Dave Masten, founder and President of Masten Space Systems commented “Masten Space and XCOR are next door neighbors here in Mojave. We’ve worked together on many tactical problems over the years and our corporate cultures mesh well. Working together on something like this simply made too much sense. We can’t wait to start working with Jeff, Dan, and the XCOR team to help NASA build affordable and responsive landing platforms.”

“Our company work ethic and styles are very compatible, and with XCOR propulsion and Masten VTVL technology, we can solve problems of national interest, and I am excited about the possibilities,” said Jeff Greason, CEO and Founder of XCOR.
Andrew Nelson, Chief Operating Officer of XCOR added, “It’s a no brainer, Dave’s team is the absolute best New Space company when it comes to VTVL and autopilot unmanned operations – they demonstrated that in October by winning NASA’s lander challenge. And we feel our LOX/methane engines are unsurpassed in the trade space today by anyone. We should bring this tandem set of best in class capabilities to NASA, it just makes sense for them and for us.”

XCOR and Masten will be jointly marketing their skill sets and services to the NASA community as prime contractors, and as joint teaming partners for larger systems integrators and prime contractors servicing the NASA community.

# # # # #

Masten Space Systems is a Mojave, CA based aerospace company developing fully reusable vertical takeoff, vertical landing (VTVL) launch vehicles, rocket-related products, and engineering services. The company’s 6000 square foot production facility and 200,000 square foot testing facility is located on the Mojave Air and Space Port. The company designs and builds aerospace solutions that focus on durability, long operational lifetimes, and minimal per-flight maintenance. For more information on the company see http://masten-space.com

XCOR Aerospace is a California corporation located in Mojave, California. The company is in the business of developing and producing safe, reliable and reusable rocket powered vehicles, propulsion systems, advanced non-flammable composites and other enabling technologies for responsive private space flight, scientific missions, upper atmospheric research, and small satellite launch to low earth orbit. The Lynx is a piloted, two seat, fully reusable, liquid rocket powered vehicle that takes off and lands horizontally. The Lynx production models (designated Lynx Mark II) are designed to be robust, multi-commercial mission vehicles capable of flying to 100+ km in altitude up to four times per day. XCOR’s web address is: www.xcor.com.

Contact:
Michael Mealling
Masten Space Systems
Phone: +1-888-488-8455 x102
Email: mmealling@masten-space.com

Mike Massee
XCOR Aerospace
Phone +1-661-824-4714 x127
Email: press@xcor.com

I can’t speak for the company, but personally I’m really glad we were able to find a way to make this partnership work. I’ve got nothing but respect for the XCOR team, and have been trying to find a way to work with them for years. As Jeff said at Space Access, it’s deals like this that show that the industry is starting to grow up.

Jeff, who happens to be in a really good position to know, clears up a common misconception about the Futron Space Tourism study that I’ve seen made a lot of places (including in the comments section here):

This author, being intimately familiar with the forecast, can shed some light on that forecast. A common misconception about the study is that it forecasts the number of people who will actually fly. Instead, it forecasts the demand for space tourism based on the level of interest among the population of people with the means to pay for such flights. Realizing that demand is a separate issue (as can be seen in the delays in opening the even-larger market for suborbital space tourism.) The supply of vehicles that could serve the orbital space tourism market was dramatically affected by the Columbia accident and the concomitant policy changes. At the time the study was released the ISS was scheduled for completion by mid-decade, with an increase in the number of Soyuz flights, all of which could accommodate at least one commercial passenger. Delays in assembling the station and the planned retirement of the shuttle have all affected the number of seats available, and have kept prices high.

His next paragraphs also make the point that Space Tourism isn’t the only commercial human spaceflight market.

I’m not opposed to other markets, and in some ways have become a much bigger fan of propellants as an early RLV market, but I think the “case” against the commercial spaceflight market has always been a rather weak one.