Saturday, December 20, 2008

Hypothetical Oxyhydrogen Propulsion Systems

Oxyhydrogen (LOX/LH2) is generally a very expensive system, but despite its overall impracticality, has an awful lot of experience in space because of the performance-oriented nature of space launch vehicles to date. Unfortunately, that doesn't make these systems practical for low-cost space access...until you are contemplating very large boosters and payloads, such as the Ascent Roadmap's Saturn-class Freezerburn. After Columbia, after almost six years of tinkering with booster concepts, has concluded the following about oxyhydrogen's proper use in low-cost boosters:

- A point will be reached where the reduction in mass of the upper stage for a given payload performance will outweigh the difficulty of increasing base area or total thrust pressure in the first stage. This is the point where oxyhydrogen becomes practical in upper stages. There is no analysis yet to back up this figure, but I currently estimate that it is somewhere between 100 and 500 tonnes LEO. Beyond LEO, oxyhydrogen may be welcome at a lower performance point. Let us assume for a moment that 200 tonnes LEO is the largest payload an all-oxykerosene launch vehicle can launch to LEO. By upgrading this launch vehicle (with the same lower stage system) to an oxyhydrogen system, a 300 tonne LEO performance can be achieved in a vehicle which is larger, but perhaps weighs less, since oxyhydrogen has a higher specific impulse. However, if one is looking for a performance beyond LEO (say, to geosynchronous) an oxyhydrogen transfer stage for higher energies may be desirable before oxyhydrogen is welcome in the booster itself. It is unlikely that this point in the Ascent Roadmap will be down where oxyhydrogen stages are currently used.
- Oxyhydrogen will never be practical in lower stages because it produces a lower total thrust pressure than other propellant combinations, including oxykerosene, which is currently the cheapest one.
- Oxyhydrogen will be best looked after by permanent launch site facilities. When selecting launch sites for oxykerosene-dominant booster, it will be important to make available the real estate that will eventually be required for oxyhydrogen. Since launch site real estate is dominated by the hazard area, and facilities can be inside it (preferably near its edge where the facilities can be practically armored against a booster failure), this should not be a problem.

The actual technical issues of using liquid hydrogen are rather unique. Some of its properties are nettlesome, but some are beneficial. Liquid hydrogen creates some problems while solving others, leading to what were once highly novel approaches to engine design. The Pratt&Whitney RL-10 is the first expander-cycle engine, one which is able to use a turbine that does not require combustion for its energy. This cycle is also being used in ESA's new Vinci engine for the ECA-B Ariane 5 upper stage. The weird properties of liquid hydrogen are listed below:

- It has a density of just 0.071kg/m3, one fourteenth that of water. This requires very light tanks with large volumes, but the low density means that these tanks usually do not require structures to handle propellant inertia or weight resulting from accelleration (such things as trusses and slosh baffles.) Their loaded vibration modes are similar to their empty vibration modes: in plain language, a tank loaded with liquid hydrogen can barely tell that it isn't empty. A disadvantage for turbopumps is that they need to generate fourteen times as much “head” as an equivalent water pump. A pump doesn't care as much about the discharge pressure of the fluid it is pumping, as much as it does its head, the product of acceleration and density. This is better known in earthly terms (since pumps have been around several times as long as rocket engines using them have been) as the column of fluid that a pump can raise against Earth's gravity. This column needs to be fourteen times as high to achieve the same pressure as for water, making hydrogen pump design a major challenge.
- It has a boiling point of just 25K, or -248degC...very, very cold. Liquid oxygen boils at a comparatively balmy 90K, and freezes well above hydrogen's boiling point. This makes materials selection and insulation a hazzle to make up for the tanks thinking that they are nearly empty.
- Hydrogen has the ability to form hydrides with many metals, very useful if you are contemplating a metal hydride battery, but very nettlesome if you're trying to build a titanium alloy tank. Most hydrides are far weaker and more brittle than the metals they are based on, and thus the tank is weakened and will fail. A nuclear reactor in Pickering, Ontario suffered enough weakening in one of its pressure tubes that it blew in August 1983 long before it was due for replacement. This was due to a small amount of hydrogen (actually, deuterium, the two-nucleon isotope of hydrogen) being dissociated from the cooling water by the reactor's radiation. This obscure accident is the closest a Canadian nuclear power reactor has ever gotten to a meltdown. Hydrogen embrittlement of pipe and pressure-vessel materials is a much bigger problem in sour gas wells and pipelines, where the metal can take hydrogen from hydrogen sulfide, which leads to both embrittlement and nettlesome sulfur deposits. Fortunately, all this experience means that while embrittlement is a nettlesome problem, it is one we are good at solving and mitigating in industries less expensive than that of space launch vehicles.
- Hydrogen has a remarkably high heat capacity. Both the specific and latent heat capacities hydrogen are far higher than those of most other substances, and may in fact be the highest of all substances. This makes hydrogen a very good coolant (although much to be desired as a refrigerant.) It is also these capacities that make possible the expander cycle. In the expander-cycle engine, the liquid hydrogen is first pumped to high pressures by the engine's turbopump (which is also pumping much denser liquid oxygen), and then fed into the nozzle jacket. Next to the main fire of the engine, the hydrogen is able to pick up a tremendous amount of heat while fulfilling its function as a coolant, keeping the engine's rocket motor from melting. This heat picked up from the nozzle is usable as enthalpy in the turbopump's turbine, and there is plenty to go around. After driving the turbine, there is still plenty of residual pressure, and the hydrogen is then injected into the engine, which can run at pressures as high as 5MPa before anything needs to be burned with the hydrogen to provide more power in the turbopump. This nozzle jacket enthalpy also makes staged-combustion engines like the SSME and RD-0120 much more efficient. Large engines have less nozzle area per unit flow of hydrogen, and so make it harder to stuff this usable enthalpy into the hydrogen, so there is a size limit to expander cycle engines with single motors. This limit could be exceeded by using multiple motors in the engine, or by convoluting the combustion chamber to have more area available to transfer heat into the hydrogen. Neither of these ideas have been tried as of yet (Ascent Roadmap is not contemplating oxyhydrogen engines large enough to bother.) The SSME and RD-0120, desiring 20MPa chamber pressures, add additional energy by burning a small proportion of the oxygen with the entire flow of hydrogen. The RS-68 is the laughing stock of oxyhydrogen engines for three reasons. First, the engine is too large to use the expander cycle effectively, and second, more importantly, the nozzle is ablatively cooled, so that transferring nozzle waste heat into usable enthalpy in the hydrogen is not even attempted. Finally, and most importantly, the RS-68 uses an open gas cycle, thus not using the entire flow of hydrogen to efficiently drive the turbine for the turbopump, but only a tiny flow of less efficient combustion gas, which is then lost to the engine by being vented out of a separate low pressure nozzle. This results in an abysmal specific impulse 9% lower than its competitors. This propagates into the almost totally ignored Delta IV launch vehicle (which piles on the bad decision of using oxyhydrogen in a first stage in the first place.)

So...what would the Ascent Roadmap ever do with oxyhydrogen? In order to consider it, we must look at the sources of expense in existing oxyhydrogen engines and seek to eliminate or reduce them. Ones that can be helped:

- Interpropellant seals: An interpropellant seal is a seal required to keep separate two fluids which can burn if they come into contact with each the rather undesirable location next to the shaft that connects pumps and turbines together. Inside a rocket turbopump, without getting into the specifics, you will almost always have a situation where the fluid being pumped will leak into the fluid driving the turbine because the fluid being pumped is at a higher pressure. The typical approach to interpropellant seals is to have a helium purge chamber and two seals: one to separate the turbine fluid from the helium, and one to separate the pumped fluid from the helium. As the two fluids leak into the helium chamber, the helium is circulated and vented to keep them from building up to concentrations where they can burn. The interpropellant seal almost never appears in industrial applications, but almost always appears in rocket engines. Life would be a lot easier if you could eliminate it. Let's see if we can.
- Diverse pumps: The properties of liquid oxygen and liquid hydrogen are so different that it is nearly impossible to achieve much commonality between the pumps and still have an engine of sufficient performance. Oxygen, with its much higher density, can be delivered by a pressure-fed system and still perform well. Especially since the hydrogen tank will be at least two and a half times as large as the oxygen tank in a typical oxyhydrogen stage, it may be possible to implement an engine that pumps only a single propellant.

So...can we design an upper stage that pressure-feeds its liquid oxygen, and uses a turbopump for its liquid hydrogen and thus eliminate these problems. Doesn't seem likely because the increased mass of the liquid oxygen tank at high pressures will make the stage too heavy. In this scenario, we would have to select a fairly low chamber pressure of perhaps 2MPa. The engine would use the expander cycle to drive its hydrogen turbopump, thus pure hydrogen will be in the turbine, with pure hydrogen in the pump at a slightly higher pressure. Since a small leakage of pumped hydrogen into the pure hydrogen turbine gas is acceptable, the interpropellant seal is eliminated.

What if we want to pump both propellants? The problem here of course is that in a typical staged or expander cycle engine, you will have a hydrogen turbine coupled to an oxygen pump, thus again requiring the interpropellant seal. Perhaps, though, since almost all of the pumping power required by an oxyhydrogen engine goes into the hydrogen turbopump, perhaps a less efficient cycle not requiring an interpropellant seal can be used for the oxygen pump. One possibility is to use a separate supply of hydrogen peroxide, as has been done in the Viking and RD-107/8 engines (famous for being economical in their time, on board the Ariane 1-4 and Soyuz boosters respectively.) Hydrogen peroxide breaks down into water vapor and gaseous oxygen vigorously enough to drive a turbine, but it has the disadvantage of requiring a third propellant to be carried on the stage (grant that it's relatively small, and the stage will probably already be using it for its reaction control thrusters, this is only a small problem.) As for the pumped liquid oxygen leaking into the turbine driven by hydrogen peroxide, the real problem is that the water vapor will freeze. This problem is bigger if it is the turbine gas that will leak into the pump. Rather than requiring a full blown interpropellant seal, the shaft can be packed with a dessicant (which removes the water vapor) and a heater (which keeps it from freezing.) Such a scheme may be overloaded, making the design of such a seal as much of a pain as an helium purged interpropellant seal, so it would help to find a way to drive the turbine with a cleaner gas mixture. The simplest solution is to burn a little bit of the hydrogen with either a generous dose of oxygen, or the entire oxygen flow. The latter is probably not practical, since oxygen does a relatively poor job of driving turbines compared to hydrogen, but don't count it out, since the RD-170 family drives their turbines with their oxygen flows burned with kerosene. The other possibility is to drive the turbine with a fraction of the oxygen flow burned with hydrogen in an open cycle (venting this oxygen rich turbine gas through another nozzle.) This turbine exhaust may also be useful as a pressurant in the oxygen tank, and of course, used for roll control during the engine's operation, saving on reaction control propellants. The turbine gas resulting from burning a little hydrogen with a lot of oxygen will still have a lot of water vapor, but not as much as hydrogen peroxide...enough to make a difference in the design of the oxygen pump's seal? We don't have to know, but we should save the pages in the recipe book, just in case.

Tuesday, December 16, 2008

Prochron Heads Up

I am currently doing a huge batch of much modernized AFAL ( - under 300k...don't bother unless you are, or want to become, a rocket scientist runs in DOS...even in XP's prompt two and white...or black and green if you find an old enough system.) The "modernization" of AFAL data involves putting it into an OpenDocuments spreadsheet and adding conversion formulas for specific impulses (to m/s), pressures (to kPa), and for ambient pressures and total thrust pressures (this latter needed to size motors for vehicle concepts).

There is some bad news for Prochron...the Isp's used in the spreadsheet and previous posts are too high for the first stage. I saw this when I dug up my April 2007 notes for Symtex. The new Isps are about 2100m/s at sea level, 2700m/s in vacuum, but it should still be possible to achieve 3000m/s in the upper stages and on the OpenLuna lander versions.

AFAL batch work is still needed for above Prochron/Symtex oxykerosene work, increased exit pressures for Prochron/Symtex, oxy/LNG and possibly oxyhydrogen for Freezerburn. I'll only do the oxyhydrogen if I get writer's block, or once I've worked through to Lilmax. The "flip point" for oxyhydrogen starting to benefit the Ascent Roadmap is probably in the 100-500 tonne LEO payload range...and my personal "point" guess is 250 tonnes before I start using it prior to achieving low energy orbit. A lot of more worthwhile work remains before I start looking for that point with calculators and spreadsheets.

Friday, December 12, 2008

Ascent Roadmap Propellant Decisions

The Ascent Roadmap choice of propellants is currently (and tentatively) liquid oxygen and kerosene for all lower stages of all Ascent Roadmap concepts, and liquid oxygen and either liquid ethylene or liquid natural gas as the propellants for upper stages. These propellants were chosen by a process of elimination which was not documented until now. Propellants were eliminated in the following order, even before After Columbia Project had decided to work on its own detailed launch vehicle concepts in 2005. Part of the reason for that particular decision was the discovery that accurately modeling a commercial launch vehicle for simulation purposes without the help of the vendor is as difficult as starting a basic concept from first principles and modeling that instead.

Solid Propellants: It really is a no-brainer, but the reasons I've had for shunning solid propellants have varied over the years. First, I thought that it was impossible to make a non-detonable solid propellant with enough performance. This probably took a circuitous route from the Bureau of Alcohol, Tobacco, Firearms and Explosives in the United States. This myth was first shattered by learning from NASA sources through egroup discussions that the Shuttle's SRB propellant is not detonable under any circumstances. It just burns. It is also very difficult to get it to burn vigorously enough to be a rocket propellant; it takes a four-stage igniter, itself generating a thrust pulse of 27kN, just to start the SRB. Standard issue model rocket propellant is not quite as benign, because it is easier to get it to burn vigorously, but it is still non-detonable. It has been tested with blasting caps and detonation cord, which can't even ignite the stuff! But, the Bureau of Alcohol, Tobacco, Firearms, and Explosives in the United States has set itself against the amateur rocket community, and has very inaccurately opined model rocket fuel as an explosive, and very inappropriately classified it as one. Still, ground safety is an important reason why solid propellants suck for large launch vehicles. The reason why is because your rocket motor comes with its full load of propellant, and weighs as much (more if it has remove-before-flight (RBF) items attached to it) as it will at lift-off. This isn't a big deal until complete rocket motors start weighing about 25kg or 50 pounds or more. This is the threshold where you can start wrecking your back in lifting operations even with the proper technique (some people, like me, using proper technique, can lift up to 100lb without any risk of long as I'm wearing safety toes.) Some further territory can be gained by dollies, pallet jacks, forklifts, cranes, bobcats, zoom booms, etc. (“MHE”), but gas pressurized transfer and transfer pumps get cheap fast compared to all this fancy stuff. Once you're into the Kilder scale, where you need all this fancy stuff anyway, it is a heck of a lot cheaper, if it need only lift some 8-10% of the lift-off weight of a stage, and the only advantage to having solids is that it saves you from having the pumps and tanks for handling liquid propellants at the launch site. Also, with solids, the propellant still has to be loaded somewhere, and mixing the non-detonable propellant mixture of tire rubber, aluminum dust, and ammonium perchlorate oxidizer is still quite dangerous. This is why the National Association of Rocketry frowns on rocketeers mixing their own propellants and making their own motors (which are small pressure vessels with plenty of complications in their own rights. If they are done poorly, they can be very dangerous.) Not having to do this at the pad is the big advantage of solid propellants over liquids (which are actually mixed during the motors operation, not prior.)

In flight, solid motors really suck. Once started, the only way to make them stop is to blow them apart. In the Shuttle program, this has actually been done once. During the ascent of STS-33, the final mission of Challenger (better known by the manifest number 51-L), the guidance system lost communications with the solid boosters, and then about half a second later, lost electrical power. These were natural side-effects of the vehicle breaking up. Both solid rocket boosters survived the breakup of the Challenger and the burnt-orange colored External Tank hanging from her belly. The right hand solid rocket motor had an estimated thirty inch diameter hole stemming from a leak in her aft field joint, but what is less well known is that another one was just opening up in its forward field joint after it broke away from the disintegrating External Tank. Of course, this difficulty in segmenting and the more catastrophic consequences of leaks is yet another nail in the coffin of large solid-fuelled motors. As the two solid boosters separated from the Shuttle, they had no guidance and in an aerodynamic sense, are statically unstable, making their flight totally unpredictable. The left hand solid booster was approaching the edge of the flight corridor when the range safety officer hit the switch. This sent a radio signal to both boosters and the External Tank (which had already broken up) to set off a set of linear shaped charges designed to cut them into pieces. By this point, the External Tank's radio receiver, battery, and explosive charges were several hundred feet apart, floating separately into the Atlantic, so it did not respond to the signal (the External Tank's charges were later recovered, removing any doubt that they caused the disaster.) The intact solid fueled boosters obediently exploded, reducing their chamber pressures enough that they quit burning. Pieces were recovered with the last ten seconds of fuel still attached. There was initially some argument as to whether shutting down a solid motor could be done safely, and in fact, Shuttle's solid boosters had been designed with such a system. The implementation was having two hatch-like blow-off panels near the nose of each booster, positioned to direct the rocket exhaust away from the Orbiter. These had to be far larger than the throat of the nozzle to vent enough exhaust to shut the motor down. When they studied what the resulting reverse thrust forces would to the aborting Shuttle stack, they gave up. Imagine if your vehicle had the brakes and tires to actually stop on a dime from highway speeds. What would actually happen is that your wheel assemblies really would stop on a dime while your vehicle's body, with you still in it, would bust loose and skid, flip, roll, and break up all over the next couple hundred feet of the highway. Even if your whole car could handle such a stop, you'd be a splat mark all over your windshield and dashboard, even with a seatbelt and airbag. This was the sort of problem the engineers were facing. They were stuck with solid boosters already, that choice having been made in ignorance (or defiance) of all the stupidities of solid motors. On a positive note, solid motors are excellent for military applications. Typically, these sit in a stockpile or silo for years, must be ready in minutes (seconds for the ones on the battlefield), and must then work perfectly when the trigger is pulled. They can also be made strong enough to survive the violence of shutting down a solid motor. The Minuteman missile has this facility in the third stage, to prevent it from flying too fast and overshooting its target area. Most satellites do not have, nor can acquire, the strength required to survive a solid shutting down, so solid satellite boosters keep from overshooting by either fishtailing to waste energy, or by having a liquid propellant final stage which is easy to shut down (and often, restart for further maneuvers.) If the final stage is a spin-stabilized solid, which can't fishtail to adjust its performance, it is built to very tight tolerances, aimed very accurately by the last guided stage, and the result is that the cost belies its simplicity, much like a $100 ball point pen (a $0.10 ball point pen (with a lid) actually has one more moving part than a typical spinning solid motor stage!! A ball point pen with a retractable tip has three more. The only moving part on a spinning solid motor stage is the valve for the thruster that keeps it from wobbling.)

The exception to these considerations for the Ascent Roadmap is very reasonable: the first stage of the Prochron. One module of this stage has a mass of just 6.6kg, well under the lifting hazard threshold, and the advantage of being able to grab a mass-produced, off-the-shelf, mostly reusable piece ofuire equipment for this role that can be used with very little preparation and no modification is a major bonus.

Solid Fuel: That meaning hybrids with a liquid oxidizer and solid fuel. Operationally, these hybrid motors have most of the disadvantages of the solid motor, and most of the disadvantages of the liquid propellants, along with some advantages. First, hybrid motors can be throttled, shut down safely, and even restarted (although not as easily as a liquid-fueled motor.) The main disadvantage that they inherit from solids is that they are still heavy, perhaps half that of an equivalent all-solid motor. The main disadvantage that they inherit from liquid-fueled motors is that the oxidizer (sometimes more than one) still has to be loaded into the vehicle. Often, a separate oxidizer is used for ignition, in addition to the flight oxidizer. The second disadvantage they inherit from liquid systems is the relative volatility compared to solid motors. Yes, liquid fueled stages actually blow up more easily than solid ones, all other factors being equal (i.e.: if you don't have leaky rotating double bore tang-and-clevis joint designs in your solid motors that can blow up school teachers on cold days.) This is the main argument for the Ares I crewed launch vehicle NASA is developing. The counterargument for the Atlas V is simply that that booster family has flown over 80 successful launches in a row, proving that the design is solid and the people running the system are not screwing up. The Shuttle has a similar record, but the Ares I is nothing like the Shuttle, and half the people running it will be laid off for about five years between the two. Even those who come back will not be current on any of the systems after a five year break. Even without the five year break, the differences between the Shuttle and Ares I will guarantee that the launch teams won't be as comfortable as confident as those running the Atlas V, and they will be more likely to screw up, even after all that expensive retraining. And of course, rockets are not very forgiving when you screw up, solids are even less so (if you drop the External Tank from the crane, you'll be lamenting the loss of the tank...if you drop an SRB segment, you'll be lamenting the loss of whatever it lands on!)

Liquid Hydrogen: It was easy to rule out liquid hydrogen...its low density (one fourteenth that of water) dictates the need for a tank that is very light for its size, and its low boiling point 250 degrees Centigrade below that of water dictates extreme materials and good insulation. The need for very light tanks dictates a turbopump system, and pumps care about head, not pressure. A hydrogen pump needs to develop fourteen times as much head to generate the same pressure as a water pump, as well as fourteen times as much flow per minute to deliver the same mass. These difficulties don't add up...they multiply. The result is expense far beyond any benefits to be had from improved performance. If you are looking for cheap access to space, don't look to hydrogen. Also, most of these arguments also apply to the “hydrogen economy” envisioned by many environmentalists. Since private industry has more sense than NASA, I'm willing to place bets that the hydrogen economy will never develop where vehicles have to go any substantial distance from the power grid (it might appear in trolley cars and metro trains, allowing some freedom for these vehicles...if propane is not allowed in underground parking, hydrogen never will be, making it a poor choice even for city cars.)

Toxic Hypergols (full names nitrogen tetroxide, nitric acid, hydrazine, mono- and dimethyl hydrazine): These used to be the favorite choice in general, but they have environmental and ground safety problems that make them unfavorable for large commercial boosters, or small ones for that matter. If applied to Prochron, the safety equipment required to handle the propellant will cost far more than the actual launch vehicle. Hypergols do have the advantage of auto-ignition, making them ideal for thrusters once already in space. For this role, the non-toxic (but still highly reactive) hydrogen peroxide is preferable if you are looking to save money. Hydrogen peroxide was ruled out of Ascent Roadmap main propulsion systems because they don't need autoignition, and because hydrogen peroxide is quite expensive to produce. It has not been ruled out for reaction control thrusters.

Nitrous Oxide, Hydrogen Peroxide: Main argument against these is the expense of production, but they are also intrinsically hazardous because they contain inherent technical terms, a positive enthalpy of formation. In simple terms, they can blow up on their own...both have a history of doing rocket propelled vehicles. The most recent such accident was when a SpaceShipTwo propulsion system exploded on the stand in Mojave in 2006, killing three people. Nitrous Oxide is hardly favorable for its autoignition properties, and is famous for its insidious effects on the nervous system (making leaks that much more hazardous) so it has been ruled out for all Ascent Roadmap systems. Hydrogen Peroxide, however, is an excellent reaction control system monopropellant, so it may yet find itself on Ascent Roadmap boosters.

Propane: Propane was briefly considered at the recommendation of MicroLauncher competitor Charles Pooley (who has missed the mark on many more important topics, such as the specifics of the microspace revolution which is now passing him by, and the difficulty of engineering pico-scale interplanetary spacecraft.) The reason why propane was considered, and then only for the Prochron booster, was its vapor pressure. Its volatility means that it will clean up its own spills by boiling away, and its vapor pressure may be usable for feeding it as a propellant. Its inherent vapor pressure is a disadvantage on any scale larger than a single Prochron modules (which is about the same size as the MicroLauncher first stage) because it creates a large, flammable and invisible cloud of gas as it boils off, one that can also displace the oxygen from the air around launch crew workers, killing them even without a spark. Kerosene has a high flash point, and for this and other reasons, a ground covering is highly desirable anyway. Spilled fuel is not necessarily lost under these circumstances. A concrete or firebrick pad (probably both) protects the soil beneath, and, if the pad is not large enough, or if the rocket is being launched from a stand without a pad, a tarp will be used to protect what will almost certainly be a grassy fuel from both fuel spills and burning debris from the Prochron first stage (the only stage in the whole Ascent Roadmap likely to produce any without a failure.) As far as vapor pressurization goes, propane in thermal contact with liquid oxygen through a common bulkhead will not develop enough pressure to keep up with the liquid oxygen. This complicates the pressurization system to the point where kerosene is just as good as propane. With either fuel, either a supplemental pressurant supply is needed for the fuel, or the liquid oxygen tank needs to be vented to bring its pressure down to parity with the fuel tank.

What's currently under consideration:

Liquid Oxygen: Pennies a litre, easy to produce anywhere (on Earth at least), pressurizes itself, and has lots of experience in rocketry. The choice of liquid oxygen is actually very easy. There are only two real competitors as oxidizers for any booster, not just those for the Ascent Roadmap. Nitrogen Tetroxide, which is hazardous (and that hazard makes it expensive) and Hydrogen Peroxide, which would work, except that it is difficult and expensive to produce.

Kerosene: It is the cheapest fuel available, and it has nearly as much experience in rocketry as liquid oxygen. It is easy to store, easy to ignite, and relatively easy to pump with its relatively high density. Most of the competitors are either more volatile or very toxic, neither of which endear them to lower stage applications. All are more expensive, counting all the secondary costs of transportation, storage, transfer, and safety. All fuels (including liquid hydrogen) are currently derived from fossil fuels, so it is very likely that all of them will follow a cost trend somewhat parallel to that of energy prices in general, and petroleum products in particular.

Liquid Ethylene: Properties are very similar to those of liquid natural gas, and it may be possible to make stages that can operate with either fuel (that would be great.) Ethylene is produced either by a Fischer-Tropsh reactor or steam fractionation, making it relatively expensive compared to liquid natural gas (LNG) and especially kerosene. So why? Ethylene is backed by three Mars Sample Return studies as very likely the best fuel to produce on Mars. This is Ascent Roadmap borrowing the long view from the After Columbia Mars Direction. Liquid ethylene is not being considered for the lower stages of any Ascent Roadmap concept. Upper stages are smaller, therefore more similar to the Mars launch vehicles that will be taking off in near-vacuum in less gravity, providing useful experience for them. Being smaller also makes the all-new oxyethylene propulsion developments more affordable, since ethylene has no experience in rockets at all.

Saturday, December 6, 2008

Prochron Spreadsheet

I've decided to post online the basic optimization spreadsheet at:

The grey ones are not optimal.

Friday, November 28, 2008

Prochron Update

Prochron 0811B will be set aside at this point. The OTRAG style configuration appears to be the way to go, but it is obviously rather limited. The propulsion for Prochron and Symtex is likely to be oxykerosene blowdown pressure-fed made of ordinary steel, perhaps with some cheap hoop wrapped composite. The selection of first stage motor is still Cesaroni M4770-P V-max...this one lasts only two seconds. Everything else is open, but the assumption for the stage is that it is 10% of its full mass at burnout. The module turned out to be 151kg, sized to optimize version 3204. The ascent is 9500m/s. The numbering system works like this:

1st digit: number of modules in Stage Two
2nd digit: number of modules in Stage Three
3rd and 4th digits: 10kg units of Stage Four

Version 1000: The sounding rocket (has 4 modules in Stage One)
Total impulse is 445860Ns, lower S class and about half of the maximum allowable under FAA amateur rocketry rules. It has to carry a payload of at least 100kg in order to stay under the amateur ceiling.

Version 3104: Design point, 3kg orbital booster (10kg listed payload allows 7kg for guidance and dunnage; 4kg has been allowed for the fairing.) Gross mass is 724kg. Has 10 Stage One motors.

Version 2104: Minimal: barely makes orbit with its guidance system (700g of payload). Gross mass is 557.5kg. Has 8 Stage One motors.

Version 4104: Payload of 5kg, Gross mass is 890.2kg, has 12 Stage One motors.

Version 5204: Payload of 9.8kg, Gross mass is 1230kg, has 17 Stage One motors.

Version 6204: Payload of 11.8kg, Gross mass is 1402.8kg, has 20 Stage One motors.

Version 6210: The biggest. Payload of 13.8kg. Gross mass is 1464.8kg. Has a bigger upper stage.

Post comment or send email if you want access to the spreadsheet. I need to know your email, post like "aftercolumbia at gmail dot com" to make sure it doesn't get filtered by blogger software. Oh yeah, that's my email address if you'd rather not have any exposure in the comments section. Check out the landers shortly to be posted at


Saturday, November 22, 2008

Prochron 0811B

It's official: Prochron 0811B specs

Booster on pad: RTL fuelled: 695kg
Booster on pad: RTL unfuelled: 141kg (Stage One installed)
Booster on truck: not fuelled: 75kg (Stage One not installed)

Stage One: Cesaroni M4770-P 6.6kg total per module; ten modules
Stage Two: Prochron module * 3
Stage Three: Prochron module *1
Stage Four: Prochron upper stage

Cesaroni M4770-P: 6.6kg rocket stage, 3.1kg at burnout, 2043m/s exhaust speed. Stages at 107m/s elapsed serial delta-v after 2.0sec. Average thrust: 4770N (47.7kN vehicle)

Prochron module: 150kg loaded, 15kg empty. Average exhaust speed 2400m/s first stage, 3000m/s further stages. Estimated 20cm diameter, 5m long (avg. density 0.75kg/L). Average thrust: 2500N (tentative; 7500N Stage Two.) Estimated burn time: 162sec (assumed constant thrust)

Prochron Upper stage: 20kg loaded, 6kg empty. Average exhaust speed 3000m/s. Estimated 20cm in diameter, 0.6m long. Average thrust: 250N Estimated burn time: 168sec (assumed constant thrust)

Fairing: 4kg, payload 5kg (i.e.: 3kg Cubesat plus 2kg deployer)

1. 3:1 Stage Two: Stage Three found optimal, needs to be verified with better optimizer
2. 6m launch tower estimated exit speed at 27m/s based on

1. Calculated Delta-v only 9258m/s due to previously undetected error in optimizer. Would have to be air launched.
2. Module burn times are too long. Try to get them under 120sec. A cost/practical requirement of the Prochron is to achieve orbit before crossing the launch site horizon.
3. Amateur/Sounding version compatibility is not known.
4. Full fledged solids in 1st stage vs. preferred hybrids. Solids are kit reloadable and made in Canada (the plus side.)
5. Does this preliminary study leave enough room for recovery systems?

Sunday, November 16, 2008

Total Thrust Pressure: Why the Big Dumb Booster Must Graduate High School

Total thrust pressure is the sum of momentum thrust pressure and exit pressure in a rocket motor. Confused yet? Well, it's rocket science. Skip to the bottom line: with only a few exceptions, the deciding dimension of the size of a liquid fuelled rocket motor is the diameter of the nozzle exit. It has to fit in under your booster, and it has to generate a certain amount of thrust to lift your booster off the ground, sea, or preceding stage of flight. As a booster designer, you also want fuel efficiency. Above all, you want a cheap ride to orbit (or else you probably wouldn't bother reading any of this.)

You want fuel efficiency? You therefore want a high expansion nozzle. The higher the expansion ratio, the more efficient the nozzle is at high altitude. There are two problems:
- On Earth's surface, atmospheric pressure tries to keep your rocket exhaust from expanding out of the nozzle. The worst case scenario is that your rocket exhaust quits expanding before it actually leaves the nozzle, separating from the walls and leaving part of the nozzle unused. This is especially annoying in a crosswind, which can cause your exhaust to separate only on one side of the nozzle (and this might happen anyway). This normally happens when a motor is first started, and is why most large boosters are bolted down until the engines are up and running.
- In the vacuum, it would be nice to have an infinite expansion ratio, but you will never have enough room in your design to do this. You therefore have to decide on an expansion ratio that will fit in your booster while also providing enough thrust. The length of the nozzle, and the skirt or truss around it to allow the stage to be carried by the booster's first stage, is generally the deciding factor, but can be mitigated in three ways. The Russians like multi-motor engines (look up the RD-0110 and RD-0124 for examples), but because not all of the base area is nozzle area, the required thrust pressure increases by at least 45.7% (assuming four motors instead of one.) The Americans like extendable nozzles (look up RL-10B-2), which slide the the bottom part of the nozzle up around the top part, much like a plastic lightsaber toy. Many providers (look up Transtage and Breeze-M) like to use donut shaped tanks (technical word is "toroidal") so that they can snuggle the upper, narrow portion of the motor up inside it. You can also see this in certain solid motors (look up Star 10).

As it relates to "big dumb boosters", I'll deal with the first stage only. As a cheap booster designer, you'll be willing to spend about the same amount of money on the first stage as the second stage, and the second stage will be smaller, and therefore somewhat more sophisticated (unless your first stage is much more reusable than your second stage, since the cost per flight is what really matters.)

I did a quick and dirty study on the effects of total thrust pressure scaling all by itself. I picked the typical parameters of a dumb booster that burns oxykerosene fuel at a mixture ratio of 2.5:1, is seven times as long as it is in diameter, uses common bulkhead propellant tanks and miraculous zero-length interstages (resulting from the desire for a quick study, not as a realistic consideration...this is why most launch vehicles, even ones matched to considerably narrower lines than their actual diameters, are considerably longer than 7 diameters...that and the fact that many of them use oxyhydrogen upper stages, which has a much lower average density.) Finally, its burnout mass is 10% of its full mass, and its payload mass is 1.5% of its liftoff mass. It lifts off with an accelleration of 12m/s2 (2.19m/s2 more than gravity, peppy for a big orbital booster, but dangerously pitiful for a model rocket.) After running it through a series of simple geometric formulae (such things as area = pi*radius squared and pressure = force/area), I came up with the following general rules about how the required total thrust pressure would scale using these assumptions:

Booster mass = 7.39 diameters cubed (diameter in m, mass in kg)
Payload mass = 0.348 diameters cubed (diameter in m, mass in kg)
Total thrust pressure required = 88.6 diameters (diameter in m, pressure in kPa)

0.1m (4 inches) diameter: 7.29kg total mass; 348g orbital payload. Requires 8.86kPa
0.2m (8 inches) diameter: 58.3kg total mass; 2.78kg orbital payload. Requires 17.72kPa
0.3m (12 inches) diameter: 197kg total mass; 9.40kg orbital payload. Requires 26.88kPa

As orbital concepts, these above are pretty fanciful. The closest real rockets are high power model rockets. They require about four times as much liftoff accelleration for safe operation, and therefore four times as much thrust pressure. Even the biggest, needing 108kPa, is fairly easy to meet.

New: 19 November 2008: Previously ignored OTRAG CRPU stage evaluated. The OTRAG system sought to achieve low cost orbital access buy building up multistage boosters from small common modules. Details at The CPRU has a loaded mass of 1500kg, diameter of 0.27m, thrust of 25kN at sea level and a length of 16m (producing a fineness ratio of ). The total thrust pressure achieved is 145.0kPa.

The vehicle failed primarily because of politics before orbit could be reached. It was being built in West Germany, and its neighbours were skittish about its potential use as a missile. This wasn't helped by the use of a Libyan test site and the theft of some CPRU hardware by the Libyans in 1983. Prime Aero (my shorthand for Boeing, Lockheed Martin, Orbital Sciences and their predecessors in the US) saw the OTRAG as a market threat. I've noticed throughout my study of history that when something happens to threaten the US military/industrial complex or energy industry, "conspiracy deception" happens. OTRAG was accused of being a cover for a joint German/South African nuclear cruise missile, and so appears to have fallen victim to this phenomenon. The primer on conspiracy deception is Nick Cook's The Hunt For Zero Point


Technically, OTRAG probably would have made orbit with only small payloads, but I'm sure, had it the chance, the company would probably have expanded to an Ascent Roadmap-like family. The CPRU, when clustered square, produces a 99.5kPa total thrust pressure, suggesting an orbital payload limit in the ballpark of 500kg.

Old Again:
0.5m (20 inches) diameter: 911kg total mass; 43.5kg orbital payload. Requires 44.3kPa
0.8m (31 inches) diameter: 3.73 tonnes total mass; 178kg orbital payload. Requires 70.9kPa
1.0m (39 inches) diameter: 7.29 tonnes total mass; 348kg orbital payload. Requires 88.6kPa
1.2m (47 inches) diameter: 12.6 tonnes total mass; 601kg orbital payload. Requires 106.3kPa

This one best matches the Orbital Pegasus booster, the one that is air dropped. It starts out at about 20m/s2, solid thrust, and masses about twice as much as this. Also, a diameter of 50in. It's actual total thrust pressure is 384kPa, hinting at an unnecessarily smart booster.

1.5m (59 inches) diameter: 24.6 tonnes total mass; 1175kg orbital payload. Requires 132.9kPa
1.8m (71 inches) diameter: 42.5 tonnes total mass; 2030kg orbital payload. Requires 159.5kPa
2.0m (79 inches) diameter: 85.3 tonnes total mass; 2785kg orbital payload. Requires 177.2kPa

The actual diameter of Taurus at 73 tonnes/1380kg payload is 2.36m. It's actual total thrust pressure is 315.5kPa, again hinting at an unncecessarily smart booster. Pegasus and Taurus are solid propellant with very high pressures, resulting in heavy stages and high drag. This is why their payload proportions are smaller. They're not "smart" in the same way as conventional boosters.

2.5m (98 inches) diameter: 117 tonnes total mass; 5438kg orbital payload. Requires 221.5kPa

This line pegs the Delta II very closely. Payload, diameter (base diameter is about 4m with the strap-ons, core diameter is 2.4m) and lift off mass are surprisingly close. Delta II also happens to be the best from a $/kg to orbit perspective in LEO On The Cheap (go to to download this excellent Air Force University Press book by John R. London III.)

3.0m (118 inches) diameter: 197 tonnes total mass; 6396kg orbital payload. Requires 268.8kPa

This line is probably the limit of practical pressure-fed boosters. It is difficult to exceed 300kPa by much with pressure-fed liquid fuelled motors. With the fancy features to do so (Tridyne pressurization, composite wound tanks), turbopumps become competitive with the better educated dumb boosters. This opinion is based on a few pages of ASME B&PV Code Sec VIII/1 calculations to estimate tank mass proportions at various pressures and a few thousand pages of AFAL Isp Calculator output (also at and the fact that conventional booster's $/kg payload cost bottoms out right here.

3.5m (138 inches) diameter: 313 tonnes total; 14.9 tonne orbital payload. Requires 310.1kPa

This line pegs the Atlas V non-strapped configurations almost perfectly, except that the payloads are smaller, at about 8 tonnes (and the actual core diameter is 3.81m.) This is probably a side-effect of the two motor arrangement, which is the least efficient way to pack multiple motors under the base of a booster (note: the Atlas V uses and RD-180 engine, which has two RD-190 class motors ganged to a single set of turbopumps. The RD-190 engine (used in Angara) has one such motor, while the RD-170 (Energia strap-on) and RD-171 (Zenit...real name for Sea Launch and Land Launch) have four such motors. American ganged engines include the MA-5 Boost (Atlas to version IIAS) and LR-91 (Titan II, III, and IV core stages). This line also pegs the Soyuz, which is an interesting study in total thrust pressure economics. The gracefully tapered shape of the Soyuz booster results from the relatively low total trust pressure of the four motor ganged engines used in the first and second stages (RD-107 used in the first stage, that is, the strap-ons and the very similar RD-108 engine used in the core second stage.)

4.0m (157 inches) diameter: 467 tonnes total; 22.3 tonnes orbital payload. Requires 354.4kPa

Here we have just passed the largest operational solid rocket motor: the RSRM. Translation for rocketry noobs: the Shuttle strap-on booster, the one that blew up the teacher in 1986. This line pegs almost every booster in the payload class: Ares I, Atlas V, Zenit, Proton, and Ariane 5. Ascent Roadmap's Lilmax is also looking for real estate on this line.

4.5m (177 inches) diameter: 664 tonnes total; 31.7 tonnes orbital payload. Requires 398.7kPa
5.0m (197 inches) diameter: 911 tonnes total; 43.5 tonnes orbital payload. Requires 443.0kPa
6.0m (236 inches) diameter: 1575 tonne total; 75.2 tonnes orbital payload. Requires 531.6kPa

An astute rocket fan will notice that both the total mass, payload mass have passed that of the 6.5m diameter Saturn IB (590 tonnes, 18.6 tonnes orbital payload). Ouchy, no? Actually, the Soyuz is about that diameter across its base. This shows that a wide tapered or squat launch vehicle is another approach to low cost space access by keeping the total thrust pressure requirement down.

7.0 m (276 inches) diameter: 2500 tonnes total, 119.4 tonnes orbital. Requires 620.2kPa
8.0 m (315 inches) diameter: 3732 tonnes total, 178.2 tonnes orbital. Requires 708.8kPa

There would be few matches past this point because the thrust pressure requirements are getting too high for any economical vehicle to match them. For example, the Saturn V (<140 tonnes orbital, 2800 tonnes total, 10.1m diameter) matches the ~7.2m line except for diameter, which lowers the total thrust pressure requirement by about 29% to 440kPa. Saturn V's actual total thrust pressure is 425kPa...pretty close for a quick and dirty study.

9.0 m (354 inches) diameter: 5314 tonnes total, 254 tonnes orbital. Requires 797.4kPa

This is Freezerburn's core diameter. It has no hope of meeting that total thrust pressure with clustered Lilmax modules. The result is that it will have a lot more base real estate than this quick and dirty study thinks such a vehicle should have, following the trend of many successful conventional boosters.

10.0m (394 inches) diameter: 7290 tonnes total, 348 tonnes orbital. Requires 886kPa

And that's as far as my quick and dirty total thrust pressure study went. It is clear that booster designers will have to take the best that today's technology can offer (staged combustion, 20+MPa chamber pressures, composite construction, preheated pressurization systems, etc.) and arrange them very creatively to launch 350 tonne payloads economically. These payloads will be asking for diameters of 10.0m or even more...we can expect their boosters to be even wider at ground level!!

Saturday, November 15, 2008

Introduction to The Ascent Roadmap

The Ascent Roadmap is a product of the After Columbia Project ( and It is intended to chart the most economical booster concepts for each class of payload

The Ascent Roadmap is also a significant part of the OpenLuna Wiki (which will eventually begin to concentrate on the Lunar Access Plan.) The following is from memory, and should be very similar in jist to what you will find at but not verbatim.

First, I highly recommend the book by John R. London III, LEO On The Cheap hosted at It has only one glitch, as does almost all advocates of Minimum Cost Design (MCD) boosters: it has underestimated the scaling problem of total thrust pressure, which makes it very difficult to design large pressure-fed boosters with payloads above 20 tonnes to Low Energy Orbit.

Payload Class Boosters:

Prochron: 3kg to Low Earth Orbit.
Earth-to-Orbit Centerpiece of: Hobby Access Plan (status=stub)
Stage One: off the shelf high power rocket motors of N impulse class. These motors provide the initial thrust to lift the booster out of the launch tower and give it enough speed so that it's passively stable design has enough speed to avoid excessive weathervaning. They are jettisoned after six seconds, and are recovered the same way as high power rockets are always (sic) recovered: parachute and puppy dog.
Stage Two (ver. 0811A): A core stage with fins, a single motor, oxyfuel (probably oxypropane) propellant combination. Additional stages, if applicable, are stacked on top. The entire assembled vehicle fits under a total impulse class of T, allowing it to be operated as an amateur long as it stays under 150km peak altitude.
Stage Two (ver. 0811B): 3 common modules, the sort inspired by OTRAG, strapped to a core stage. Each module is a blowdown single motor stage (probably oxypropane). This means that Prochron amateur rockets could use this module, and possibly be far smaller than ver. 0811A.
Stage Three (ver. 0811A): a single common diameter stage stacked on top of the second stage core, with a small pressure-fed oxyfuel motor, smaller than that in Stage Two.
Stage Three (ver. 0811B): The single module in between the Stage Two modules, carrier of Stage Four (if applicable)
Stage Four: a small upper stage designed to apply a precise impulse bit to the payload to get it into the proper orbit and correct for errors in earlier stages. It carries a small cold gas RCS system.

Operation: Prochron flies passively (or semi-passively) through Stages One and Two, which operate in the atmosphere. Stages Three and Four require active guidance and stabilization to correct from Stages One and Two, and because without air, fins don't work. The fairing is jettisoned during Stage Three. The payload is based on CalPoly's Cubesat ( 3kg forms a single "pod" of CubeSats.

Symtex: Up to 500kg to orbit
Earth-to-Orbit Centerpiece of: nothing so far
A modular pressure-fed launch vehicle intended to fill the niche underneath the Orbital Sciences Pegasus and Minotaur class boosters. If it entered service as a commercial launch vehicle, it would have no competition, except for the very largest configuration, which nearly matches Pegasus/Minotaur performance.

Common Motor: a single oxyfuel (probably oxypropane, possibly same as Prochron Stage Two ver. 0811A) used by all stages except for the upper stage, which mounts a smaller Space Motor (probably same as Prochron Stage Three).

Strap-on Module: Wide enough only to mount one Common Motor, this small module provides high lift-off thrust for about 40 seconds.

Core Modules: Mounts any number of motors from one to four (including the possibility of a single high expansion version for a second stage.) They will probably be cut to two different lengths. Try not to think of them as putty-like: stretching and cutting stages is not so simple.

Upper Stage: Contains the guidance system, is multiple maneuver capable, and contains the payload interface and vehicle electrical sources. A special type of internal gear expander coupled pump (think "displacement turbopump") is being considered. It is based on the StarRotor design (, and StarRotor Corporation has been queried for its potential use. Special thanks to them for providing After Columbia Project with basic application formulae and the University of Idaho ALLPROPS software.

Operation: Symtex is actively guided all the way to orbit, but it is intended that it be passively stable until at least the first staging event. It therefore does not require a tower or rail for weathervaning control. For conventional launch vehicles, typically extensive balloons and often patrol aircraft are used to measure winds aloft. To avoid the expense, Symtex will provide extra margin for corrections, and extra thrust to get out of the atmosphere quickly, and thus mitigate the performance effects of wind shear. The smallest configurations use strap-configured core stages, who's core operates primarily as a second stage. The largest configurations will probably use Long Cores strapped by Short Cores, plus a final Short Core as a second stage. The upper stage will probably be used in such configurations for circularization maneuvers.

Kilder: 1000 to 8000kg Low Energy Orbit
Earth-to-Orbit Centerpiece of: Lunar Access Plan

Kilder features an pressure-fed oxyfuel (probably oxykerosene) core which must be strapped in order to rise from the pad. The strap-on system provides the first stage, and a squat, rotorpump engined (probably oxyethylene) upper stage comprises the third stage. The most advanced of the strap-on systems carry Lilmax turbopump engines, once that engine has been developed for Lilmax.

Lilmax: 15 to 50 tonnes Low Energy Orbit
Earth-to-Orbit Centerpiece of: Mars Access Plan
Lilmax begins to approach the size and complexity of typical boosters in terms of its systems and launch configurations, even though it is made of spun/rolled steel. Compared to extant boosters, it considers some very unusual features for its lower stage module, along with a more conventional upper stage. The modules come in a single length, without stretching or cutting down, however the upper stage will probably have two different sizes to accomodate different payload levels and final orbits. Unusual features (unusual both to "big dumb boosters" and conventional launch vehicles) are:
- Segmented tank construction: This allows rail and truck transportability of individual segments, which form a much larger stage when assembled. This is currently done with the Shuttle's solid boosters.
- Crossfeeding of propellants from strap-ons to core, allowing the core to remain full until the strap-ons burn out and are jettisoned.
- Pressure-stabilized structure. The normal tank pressure is 200psig with a burst factor of safety of 2.0, so it may be able to stand on its own anyway. If it can't, we'd rather use cradles and removable internal webs during transport than add stiffeners to the design. Most boosters are stiffened adequately to withstand handling loads without pressure, but require internal pressure to withstand flight loads; the SpaceX Falcon describes this as though it is unusual, but it is actually the normal case.
- Recoverable and Reusable: Whatever range Lilmax winds up with, at least the first set of strap-on modules will be reusable, even if they hit dry land. Dry land may be preferable because trucks are cheaper than ships.
- High tank pressure ensures simplified turbopump design, and large safety margins for flight loads and recovery.
- Modules will have a blowdown phase late during their operation (this is also normal for most boosters), what is different with Lilmax is that after shutdown (for recoverable modules only), the tank residuals boil off and rebuild pressure for recovery. During the blowdown phase, the engines are throttled back because less suction pressure is available for the turbopumps.

Bluestar: 8000kg LEO payload fully reusable booster and orbital service vehicle
Earth to Orbit Centerpiece of: Station Access Plan (status=stub), used to be International Space Transportation System
Bluestar has evolved considerably since I first drew it up in 2001. Since then, I have discovered that putting wings on spacecraft is about as silly as equipping submarines with screen doors and wings to allow them to fly through the air. The Space Shuttle suffers horrendous expense because its wings, control surfaces, hydraulics, and landing gear (including 100lb of tire air) are carried all the way to orbit and do absolutely nothing to add to the Shuttle's functionality in space. 60 tonnes of nothing more than ballast until less than half an hour before landing. 55% of the Shuttle's orbital mass is used only for the last 0.17% of a twelve day mission. At first, Bluestar was based on Saenger II and Ashford Spacebus ( ... from memory, yikes!) but is now based on Lilmax.

Bluestar now implements a ballistically recoverable upper stage for ascent on the Lilmax lower stage modules. For basic LEO (Low Energy Orbit) services, more than one variant may be required. The satellite payload version is described at A station logistics variant would better resemble the Sprint Crew Ferry ( where you will see why I lament not preserving Brian Heick's excellent 3D artwork of this concept in support of the Orbiter Mars Direct Project.) Inside such a ferry would be the racking one would normally see inside a Leonardo class module or a SpaceHab Logistics module.

Freezerburn: 100 tonnes and up
Centerpiece of: Mars Access Plan (Phase II)
The colonization of Mars, and possibly other destinations, will require cheap heavy lift services, a very challenging mission. Heavy lifters will also be required for space hotels, and in the unlikely event they should prove economical, solar power satellites.
- Basic Freezerburn core is formed by adding structural end pieces to four Lilmax modules. These four modules then form the basic Freezerburn core stage almost for free. This is not unprecedented at all: the Saturn IB stage was formed very similarly from eight Redstone tanks and one Jupiter tank.
- Up to 12 Lilmax modules can be strapped onto the core. At least four would normally be strapped onto the core, because with only the core, Freezerburn's capacity would not exceed that of the largest Lilmax configuration, which uses five modules.
- A very large, 9m diameter upper stage is used on Lilmax. It is not recoverable, but may be reusable by incorporating it into structures in space. The most likely propellant combination for this upper stage is oxymethane, or LOX/LNG (LNG being 16% ethane, 84% methane, with trace amounts of hydrogen sulphide and helium.)
- The most likely growth option is to extend the upper stage into a booster core which sits on top of the four module core already in place. On to this, Lilmax strap-on modules above their equivalents strapped to the first stage would be attached to the upper stage. Intermediate is to use the widebody core directly on the first stage with an older upper stage on top of it.

This concludes the Ascent Roadmap introduction!!