Monday, December 20, 2010

Feasibility of a Manned Mars Exploration Mission Concept

(see also 1-11-2011 post for an update to this study)


A manned mission to Mars was investigated for feasibility. The objective was significant exploration, not a single Apollo—style “stunt” landing. It seemed insane to go to all the trouble and expense of sending men to Mars, and not visit several different sites. It also seemed insane to launch so much equipment and not try to reuse it. This consideration eliminates very small vehicle designs.

Earth orbit rendezvous and Mars orbit rendezvous were combined into a single mission design to save weight. Unmanned assets were sent by Hohmann transfer ahead of a fast-trip manned vehicle to save weight. Self-rescue or escape capability was designed-into every mission phase as much as possible. The landers and the manned vehicle were designed as reusable single-stage items.

The unmanned assets were sent single-stage one-way to Mars orbit using the lander propulsion to save weight. Lander and propellant tank assets were left in Mars orbit for refueling and re-use by subsequent missions. The manned vehicle was recovered in Earth orbit for refueling and reuse.

The manned mission time was under 1 year, eliminating the need for voluminous and heavy artificial gravity by spin (because the only requirements currently understood are 1 gee at 4 rpm). The flight deck in the habitat module was assumed to be radiation-shielded against solar flares by water and wastewater tanks, plus a little steel plate. Crew size was 6. Every component was small enough to be launched by a Falcon-9-heavy booster, or smaller. But, the interior volume is comparable to the old Skylab station, thus promising effective alleviation of long-confinement psychological issues.

All assumed propulsion was nuclear thermal rocket (NTR). The lander engines were assumed to be slight updates to the old NERVA solid core technology last tested in 1973. The manned vehicle engine was assumed to be a radiator-cooled gas core NTR, approximating a design that came within about 2 years of first article test in 1972, when all such work was stopped. This old design’s Isp was 6000 sec.

Only launch costs were estimated, as near $8 billion in today’s dollars for 16 landings during the one mission. A trade study evaluated cost reductions available for reducing landings per mission. Programmatic costs, and technology development and hardware production costs, were not estimated. It is thought that getting the gas core NTR technology ready would be the pacing schedule item for such a project. These estimates are only rough hand-calculations done by pencil and paper. They are just good enough to demonstrate feasibility, and to serve as a startpoint for more sophisticated analyses.

Basic Mission Design Approaches, Constraints, And Assumptions:

Fig. 1 (below) shows a list of the fundamental considerations. Top-of-the-list is to design-into every aspect a “way out”, meaning escape or self rescue, if in any way possible. This is the very essence of “man-rating”.

The next two items relate to flying the manned portion fast, in order to cut travel times under a year and allow deletion of the need for artificial gravity, in accordance with experience obtained on the International Space Station (ISS). Mission times exceeding a year require artificial gravity, because we have nothing but indirect (surrogate) data such as bed-rest experiments to suggest otherwise. It is unethical to risk the lives and health of astronauts on indirect data, even if they are willing. Artificial gravity by spin has to be designed for one full gee at no more than about 4 rpm, based on what we currently know by direct experiment. This leads to large, heavy, and costly vehicle designs. Choosing instead more advanced propulsion for the fast trip is very likely the easier, surer, course.

The fourth item says to do both Earth orbit rendezvous and Mars orbit rendezvous in order to save weight, plus obtain further weight savings by sending unmanned vehicles on min energy trajectories. The lander vehicles and the propellant supply supporting their operation can be sent ahead of the manned vehicle, which would then rendezvous with these supplies at Mars. For crew safety, the propellant supply for the return voyage cannot be sent this way, even though it would save considerable weight to do so. This is because the crew would be stranded, and would die, if rendezvous at Mars failed for any reason. If the manned vehicle has sufficient propellant on board to return, this outcome is avoided.

It makes little sense to go to the trouble of sending men to Mars, and not do some serious exploration. This mission study is based on doing 16 separate landings at widely-separated sites, of up to a week each, while the manned vehicle is there. This increases enormously the information return from the mission, a sort of “shotgun-pattern” planetary survey. It might even be possible to begin planting “prospecting” bases on the next mission, instead of further initial exploration.

This multiple-landing plan is subject to some safety constraints, as the last item indicates. The concept explored in this study is that 3 of the crew visit the surface, monitored from orbit by the other 3 crew. There must be at least one ready lander in orbit, in order to perform a rescue landing, if need be. This rule would terminate the mission upon a lander becoming unserviceable, unless at least three landers are sent to Mars. To minimize crew risks, any rescue landing is piloted by a single crewperson.

To make such a plan work, the landers have to be reusable, so that only three need be sent. To keep from sending lower stages, the landers must be single stage. That means they must be nuclear. As shown in fig. 2, the idea is to separate the lander propellant supply into three parts corresponding to the three landers, and send these to Mars using the lander engines themselves. These assets would remain in orbit at Mars to be refueled and reused by subsequent missions.

Fig. 1 -- Basic Mission Criteria

Fig. 2 – Sending Landers and Lander Propellant Unmanned One-Way to Mars

Fig.3 shows the thinking behind the design of the fast-trip manned vehicle. There are serious considerations for radiation sheltering during solar flare events. In the habitat module, there must be a space surrounded by water and wastewater tanks, and a little steel plate, in which the crew of 6 could shelter during a radiation storm. Prudence dictates that this be the vehicle’s command deck as well, so that critical mission maneuvers can be conducted, storm notwithstanding. The mission should take no more than about 9 months. The vehicle should be stocked with over a year’s supplies, “just in case”.

One of the requirements often soft-pedaled or ignored is volume of space available per crewmember. The psychological impact of prolonged confinement in tight spaces is a very real danger, one that can be confirmed by any prisoner who spent time in solitary. Most crew habitat designs I have seen provide about the same space as was in the Apollo capsule, which is about like a modest bedroom closet per man. That is simply not enough. The space should be comparable to that available to a family of 4 in a small (1200 square foot) house, ideally. The old Skylab space station, at 90 tons, comes pretty close to the size of habitat that is needed. It provides the baseline for this study.

Fig. 3 also shows two main engines, needed for redundancy, and a round trip propellant supply, needed in case rendezvous fails at Mars. There are two crew return capsules, each large enough to carry all 6 crew, but twinned for redundancy. These need to carried along, in case maneuver propulsion fails on the return voyage. In that event, a free return must be attempted in the capsules as the vehicle flies by Earth. These capsules must have enough delta-vee capability to “hit” an acceptable reentry corridor. That means they probably need a small propulsive service module or supply.

Fig. 3 – Safety and Design Considerations for the Manned Vehicle

Fig. 4 shows some practical launch vehicle constraints for assembly of the Mars mission in Earth orbit. Many different items could be selected, these simply correspond to a family of very cost-effective launch rockets. Data were taken directly from the Spacex website, as it exists at the time of this writing. The biggest impact is the size of objects to be launched. Payload mass is more important than fitting within the “factory stock” payload shroud. A lot of the Mars mission components could in fact ride “naked” on top of the launch rocket.

Fig. 4 – Launch Rocket Constraints on Mars Mission Component Designs

Rough Mission Delta-Vee Estimates:

Two scenarios needed investigation: a basic min-energy Hohmann transfer ellipse for the unmanned vehicles, and an “almost straight-line shot” fast-trip high-energy trajectory for the manned vehicle. Of these, determining a realistic delta-vee requirement for the fast trip is actually easier. One simply divides a representative straight-line path length (in this case about 100 million kilometers (km) by a tolerable trip time (for this analysis about 75 days). Assuming impulsive delta-vee events at each end of the trip, one obtains a nearly square-wave velocity trace vs range, because at these speeds, the sun’s gravitational deceleration or acceleration effects are negligible, as is path curvature. See fig. 5.

Fig. 5 – Fast-Trip Scenario Approximation

The average velocity over this trace is very nearly the delta-vee value required to start it, and also to end it. Thus, twice the average velocity is pretty close to the one-way delta-vee requirement for the trip. For a two-way trip, one doubles this again, to about 4 times the average velocity. For these numbers, the two-way fast trip delta-vee requirement is a very demanding 61.72 km/sec. For practical single-stage vehicles, this corresponds to specific impulse (Isp) requirements closer to 6000-7000 sec than the 900-1000 sec of a NERVA-type solid core NTR. Hence the selection of gas-core NTR technology with a waste heat radiator to effect engine cooling.

For the unmanned vehicles making a one-way trip by Hohmann ellipse transfer, the estimate is made by classical orbital mechanics methods. Calculations were made for the average orbital velocities of Earth and Mars around the sun. An ellipse was fitted between the average distances of Earth and Mars from the sun, and its perihelion and apohelion velocities calculated. Escape velocities were calculated for Earth and Mars, and circular orbit velocities calculated for a low Earth orbit (LEO) altitude of 300 km, and for a low Mars orbit (LMO) altitude of 200 km. At Earth, the delta-vee to escape on a trajectory to Mars was estimated as the difference between escape and circular velocities, added to the difference between transfer perihelion and Earth orbital velocities. The delta-vee to capture at Mars was calculated as the difference between apohelion and Mars orbital velocities, added to the difference between Mars escape and circular orbit velocities. This is probably over-conservative. See fig. 6.

Fig. 6 – Rough Estimate for Hohmann Delta-Vee Requirements

The planes of the transfer ellipse and the straight-line “shot” are more or less in the plane of the ecliptic. Thus the plane of LMO achieved this way will be inclined relative to Mars’s equator, by around 25 degrees. The landers must have some amount of plane change capability, in addition to the delta-vee necessary to land without aerobraking (in the extremely thin “air”). The surface circular orbit velocity is larger than that at 200 km, and makes a good rough estimate of minimum delta-vee. Factoring this value up by about 1.10, accounts roughly for gravity and drag losses. The plane change requirement is figured from an isosceles triangle on the surface circular orbit velocity. These velocity increments are summed for the one-way delta-vee requirement, and doubled for the two-way trip. A maximum plane change requirement of 30 degrees was assumed arbitrarily. By judicious choice from an inclined orbit, this capability brings the majority of Mars’s surface within reach of the landers. See fig. 7 below.

This 11.52 km/s value is the maximum. Not all sites require a 30 degree plane change. The minimum is no plane change at all, for the much smaller two-way delta-vee requirement of about 7.84 km/sec. Such missions have a substantially-smaller propellant “burn”. Such propellant savings, plus the very capable nuclear engine, would very likely make an orbital mission to Phobos possible during this same exploration mission, using the same equipment.

Fig. 7 – Rough-Estimate of Lander Delta-Vee Requirement

Lander Rough-Out:

The basic layout of the lander is propellant tank-as-airframe, topped by some sort of command cabin big enough for 3, and equipped with long landing legs disposed around the nuclear rocket engine. The width of the footprint should be comparable to the overall length of the vehicle for stability, so this vehicle is rather “squat” in its proportions. There should be some sort of deployable crane arm to provide a hoist to the surface. There should also be some sort of deployable solar panels to augment fuel cell electrical power. Given a heavy solid core engine and extensive landing leg structures, an inert fraction of 20% seems reasonable to assume. Combined with a propellant fraction of 70% and a payload fraction of 10%, the mass ratio is compatible with the solid core NTR Isp of 1000 sec, and the max plane change round-trip delta-vee requirement of 11.52 km/sec.

The payload comprises the crew of 3 with suits, 2 weeks of air, food, water, and fuel cell reactants (conservative for a maximum 1 week mission in case of trouble), a 3-man rover car, an inflatable Quonset hut with camping and cooking gear, and half a metric ton of scientific equipment, to include a small drill rig. Assuming the command cabin structure itself to be part of the payload, I rough-estimated 6 metric tons for payload. Thus the whole lander fully-fueled is 60 tons, with a propellant weight of 42 tons, and a dry-tank weight of 18 tons. I assumed half a ton of waste was left behind at takeoff, in making propellant usage calculations. See fig. 8. Note that an empty lander is within the payload weight to LEO of a Spacex Falcon-9 booster, although not within payload shroud dimensions.

Fig. 8 – Rough-Out Lander Design

This lander design with a 180-200 KN thrust solid core NTR engine does not have quite enough thrust to leave the surface of Mars fully fueled, but can easily take off partly-fueled, after landing from orbit. A slightly higher thrust specification would make fully fueled surface takeoff possible, but at the cost of a slightly heavier and larger engine. This is not really a necessary requirement for explorations conducted from LMO. Performance is compared to requirements in fig. 9 below.

Roughing out this vehicle is a supremely important prerequisite for the rest of the mission and vehicles because it is a major payload item, as is the crew habitat module. It should be noted that it is specifically the choice of nuclear propulsion that makes a single-stage lander possible. The delta-vee requirements for a single stage lander are simply out of the practical range of mass ratios for chemical propulsion. Without a reusable single stage lander, the mission exploration return is very much diminished: we are more-or-less back to a very few Apollo-style “stunt” landings. With chemical propulsion, a staged lander may only be used once (although an upper stage might be reused with a new lower stage). The number of landing sites is then no more than the number of landers carried to Mars, and this has a far greater effect on vehicle weights, mission complexity, and costs.

For the design and mission selected here, the average lander mission consumes some 37.56 metric tons of liquid hydrogen (LH2) nuclear rocket propellant. The plan for 16 such landings during the course of the mission then requires the delivery of some 600.89 tons of propellant as payload to Mars to support lander operations. That delivery requires even more propellant for the Hohmann transfer, even with using the lander propulsion as the propulsion sending these vehicles to Mars in order to save weight, launch costs, and complexity. See the unmanned vehicle rough sizing below.

For safety and self-rescue purposes, lander operations are envisioned as a series of sequential single landings, each with a crew of 3, while the other 3 stay in orbit to monitor progress and provide rescue capability with another ready lander. Thus at least 2 landers are required, and unless there is a third, the mission ends if one is rendered inoperative for any reason. That is why this mission plan sends three landers to Mars. Rescue is envisioned as risking only one crew as pilot in the rescue lander.

Fig. 9 – Rough Lander Performance Estimates

Crew Habitat Rough-Out:

There are three fundamental crew survival issues that must be addressed for any mission involving months of travel beyond Earth’s Van Allen Belts. These are (1) radiation protection (solar flares and cosmic rays), (2) protection from medical deterioration due to microgravity, and (3) sufficient habitat volume to stave off the psychological effects of prolonged confinement.

This mission is nominally 9 months, and certainly under 1 year in total duration. The dose of accumulated cosmic radiation is minor. The probability of a solar flare event is quite high. Therefore, there must be a zone inside the habitat shielded by water and wastewater tanks, and a little steel plate, which can support 6 crew temporarily during the event. Safety demands that this shelter also be the ship’s command deck, so that critical mission maneuvers may be flown, radiation storm or not.

The under-1-year mission time is within the realm of experience we have with microgravity exposures on the International Space Station (ISS). Therefore, this design need not provide artificial gravity by spin. The ISS exercise regimens will be adequate. This is very important, because provision of artificial gravity greatly adds to the habitat and vehicle size, weight, and complexity. The design requirements for such artificial gravity are still poorly understood, since the direct experimental work has never been done. We have only imperfect, indirect evidence from surrogate studies, such as bed rest. It is unethical to subject a crew to serious life and health risks, based on no better evidence than that. Therefore, the best design criteria we have are to provide one full gee at no more than 4 rpm. Any slower trajectory pushing total mission time beyond 1 year must deal with this design issue.

The habitat volume per crew issue is something ignored in design studies such as “Transhab”. Most of these designs would confine a crew in a space per person not much bigger than a typical bedroom closet. This is very likely to be psychologically very unhealthy, as any prisoner who has served time in cramped solitary confinement can testify. A design volume per person more like that in a lower middle class home would be far preferable. Units this large would resemble the old Skylab station in dimensions, and would be very difficult to launch. But, such a habitat could be assembled from smaller modules. It would be a part of the payload of the manned vehicle, the crew return capsules being the other part.

This study’s design is comprised of three modules, each 32 metric tons, docked in LEO to form a 96 ton habitat, stocked with substantially more than a year’s supply of food, water, oxygen, and other supplies. Such a habitat is very close to the mass of the old Skylab, but has a longer, narrower form factor. One of these modules would contain the radiation-shielded command deck. Each of these modules is within the near-term projected payload capability to LEO of the Spacex Falcon-9-heavy launch vehicle. Again, payload shroud constraints may be violated. See fig. 10.

Fig. 10 – Three-Piece Assembled Habitat Module

Crew Return Capsules:

For safety purposes, it is critical that these be carried with the habitat on the entire round-trip mission. This is because the vehicle propulsion might fail on the return trip, leaving no way to slow for capture. In that event, a crew return capsule capable of making a free return reentry at speeds very significantly higher than Earth escape speed offers the only avenue of crew escape. Each capsule should be capable of carrying the entire crew of 6, and there should be two such craft for redundancy. Some amount of service module propulsion is required to effect a proper reentry angle for survival.

Such a capsule already exists in its initial form as the Spacex Dragon. Dragon has crew capacity up to 7, and a heat shield rated for free Mars return. It fits a Falcon-9 launcher, although fitments need to be changed to accommodate some extra propulsion. I simply guessed this add-on propulsion module at 2 metric tons each. This plus the empty Dragon should be in the vicinity of 22 tons. See fig. 11 below.

Fig. 11 – Modified “Dragon” as the Crew Return Capsule, Two Required

Common Propellant Tank Module:

This is a more sophisticated design item than it first appears. There is a need to store LH2 for months at a time in zero-gee conditions. That requires what amounts to a double-shell tank, essentially a Dewar, with protection from solar thermal radiation, and at least a little meteor protection. As stackable modules, these require substantial structural strength. There is some sort of cryo-cooler (or a suitable equivalent) equipment required, plus the solar power to run that. There is considerable interconnect piping to meld these modules into an integrated propellant supply, plus a kit of extra pipe lengths and fittings to make those interconnections. It will be a substantial design challenge to achieve this in a 10% inert weight budget. Loaded tank size is set by the projected Falcon-9-heavy deliverable LEO payload weight of 32 metric tons. See Fig. 12 below.

Unmanned Vehicle Rough Sizing:

A part of the payload for this vehicle is an empty lander (18 metric tons), whose engine is also the unmanned vehicle propulsion. Inert weights of 3.2 tons per tank module add to this payload to comprise the dry-tank “burnout” weight. Loaded weights of 32 tons per tank module add to this payload weight to comprise the departure “ignition” weight. Thus vehicle mass ratio and delta-vee capability is a function of the number of propellant modules in the vehicle stack. Enough untapped modules need to arrive at Mars to support the mission’s lander operations. One third of that requirement (rounded up to the next largest number of tanks) is carried by each of the three unmanned vehicles, each with a lander. Those untapped tank modules are the remainder of the vehicle payload. See fig. 13 for a pictorial, and fig. 14 for estimated vehicle performance on its one-way mission.

Remember, after the mission concludes, the landers and empty tank assets are left docked in LMO. Subsequent missions need only bring more propellants, and reuse the landers, up to the lander engine lifetimes. Empty tank assets could be cannibalized for other purposes in future missions.

One other note: the same basic vehicle design is suitable for a variety of inner solar system missions. One simply stacks up enough common tank modules to meet the mission velocity requirements.

Fig. 12 – Common Tank Module

Fig. 13 – Unmanned Vehicle Stack (One of Three)

Fig. 14 – Rough Estimates of Unmanned Vehicle Performance

Manned Vehicle Rough-Sizing:

The design approach for this vehicle is very similar to the unmanned vehicles, only the payload and propulsion is different. The payload comprises the three-piece habitat module, plus two crew return capsules. To the radiator assembly (30 tons) and twin gas core NTR engines (half ton each of two for redundancy), one adds 3.2 tons per tank module for the inert weight total. This plus payload is the dry-tank “burnout” weight. To this, one adds 28.8 tons of propellant per tank module, to arrive at the departure “ignition” weight. As with the unmanned vehicles, mass ratio and delta-vee is a function of the number of tank modules in the stack. See fig. 15 below.

In this particular design, all the propellant required for the two-way trip is included in the vehicle. It would save weight to send the return trip propellant as a Hohmann transfer unmanned package, but, this incurs the risk that the manned vehicle might not be able to rendezvous with the unmanned fleet. In that event, the crew would be stranded, and would die. Abort scenarios where capture is avoided at Mars for an immediate return home would also be impossible. From a safety standpoint, it is simply more prudent to fuel the vehicle to be able to return home independently of all the other mission components.

Fig. 15 – Manned Vehicle Design for the Fast Trip

The Missing Technologies: Solid and Gas Core Nuclear Rockets

The key element to the fast-trip design of the manned vehicle is, of course, its engine. This is presumed to be a gas core open-cycle version of the basic nuclear thermal rocket. Unlike the NERVA-type solid core engine design of the lander, the gas core machine was never tested as a rocket engine. However, it did undergo component bench tests verifying containment of the uranium relative to the hydrogen, at about 1000:1 hydrogen:uranium flow rate ratio. It also underwent bench tests verifying controlled gas phase nuclear fission. The gas core engine was about 2 years away from a first-article rocket test when the program was shut down in 1972. The mission plans at that time allowed about 15 years to test and perfect the design, before potentially employing it on a manned Mars mission then scheduled for 1987. Most of this history is forgotten today.

The open cycle gas core NTR was thought to be adequately cooled by regenerative cooling, up to power levels corresponding to Isp around 2000-2500 sec. Above that power level, regenerative cooling was known to be inadequate. This necessitated use of a high-temperature radiator to cool the engine, whose characteristics are still guesswork. It was also thought there was a power limit above which the engine would vaporize itself due to propellant transparency to all radiation, up around 10,000 sec Isp. The planned Mars engine for 1987 was an Isp = 6000 sec design, well under that poorly-understood upper limit. That same projected design is assumed for this mission study. See fig. 16.

Fig. 16 – The Radiator-Cooled Open-Cycle Gas Core Nuclear Thermal Rocket Engine

Assuming that the developed engine and radiator system have characteristics even close to what I used for this study, then the performance of the vehicle powered by it can be calculated with at least some confidence. The results of have 6000 sec of Isp available is astounding, as given in fig. 17. Comparing this plot to the performance plot for the unmanned vehicle (of crudely similar size), one can see the difference in the delta-vee levels achievable: several tens vs only several km/sec.

Fig. 17 – Rough-Estimated Performance of the Gas-Core Manned Vehicle

As a comparison, fig. 18 below illustrates the updated NERVA engine used in the lander. That technology was substantially mature, and this shows in the quoted data in the figure. It should be noted that the lander design as worked out uses one engine, not a redundant two or three. The problem is one of thrust against gravity, and scalability of the nuclear design. The size used herein is not all that far from the original NERVA. There is some question whether a much smaller engine could even be made to go critical and produce power. Engine-out under gravity means that the remaining engines must throttle-up thrust levels to compensate for the lost engine. It might actually be easier to simply redesign the basic engine to be more reliable. This is an issue needing investigation before any designs can be finalized.

Fig. 18 -- Solid Core Lander Engine Based on NERVA Technology

Mission Information Return vs Mission Cost:

The direct launch costs are “retail”, based on number of payloads times the cost for the appropriate launcher. The data were obtained from the Spacex website, including projected costs for the yet-untested Falcon-9-heavy vehicle, and the Falcon-9 vehicle currently in flight test. On this basis, the total launch cost to LEO for 3 unmanned and one manned vehicle is right at $8 billion. That “buys” 16 landings, each up to a week long, at 16 separate and widely-dispersed sites on Mars, up to 30 degrees worth of plane change from the ecliptic. It also buys hardware that can be used again on subsequent missions, and other missions in the inner solar system.

There are hardware development and production costs to be considered, and programmatic costs. The habitat modules, the common propellant tank module, the modified “Dragon” crew return capsules, and the lander, are all items needing a “normal” amount of development, in the aggregate perhaps totaling around a billion dollars. The updated NERVA engine for the lander would actually require very little development. On the other hand, the gas core engine would be a very serious development item. Taken together, those two engines might total around a billion or two dollars. That is a wild guess predicated upon these projects being done by lean, efficient organizations. For “business-as-usual” with large, inefficient organizations, one should probably double or triple those estimates. Thus, the lower bound “wild guess” is then about $11B for 16 landings on Mars, all in one trip. See fig. 19.

Fig. 19 – Summary of Mission Design Characteristics

Reduced Mission Scope?

It is entirely possible to lower costs by reducing the number of landings under the same mission safety rules. The minimum is three. Only the three unmanned vehicles reduce in size, the manned vehicle is unchanged. But on a dollars-per-landing basis, that would be a very inefficient thing to do. See fig. 20 below for estimated savings from reducing the number of landings, total mission cost, and the prorated per-landing cost, using $11B as the total for a 16-landing mission.

Plus, there is the political effect of “getting much of the initial exploration done” in a single trip with 16 landings, in such a way as to enable future prospecting-base missions, and eventually, a colony. Compare that to the single-landing scenario, which would require more exploration missions before anything else could be done. Each and every one of these follow-up missions could be cancelled.

As a proper exploration strategy, allow me to suggest a “shotgun-pattern” planetary survey with the full 16 landings, perhaps to be followed by a second exploration mission of fewer landings at the most promising sites uncovered by the first mission. These fewer landings in the second mission would stay substantially longer times on the surface, rather similar to those proposed in “Mars Direct”.

Fig. 20 – Cost Trades vs, Number of Landings in a Single Mission

Once the exploration planetary survey is done, we are ready for a different type of mission in which “prospecting” bases are built. This is the type of mission where the in situ resources begin to be utilized, and the first indications are obtained as to what trade commodities there might be, and what the trade economy might be. These are the prerequisites for an actual colony in the future.

Alternatives to Gas Core Nuclear Thermal Rockets:

To do the fast trip manned vehicle with solid core technology requires a throwaway staged vehicle, which is neither cost effective, nor conducive to authorizing follow-on missions. Slowing to Hohmann-transfer speeds puts the total mission well over a year, which requires artificial gravity (and the resulting huge impacts on vehicle weight, size, and complexity, as well as costs). This path is not recommended.

One could spend efforts developing a flightweight nuclear electric power plant in the multi-megawatt range. Then the same manned fast trip could be done with VASIMR, or something very much like it. Developing such a power station is likely about the same risk as producing a gas core NTR engine. There would be a larger inert weight for the manned vehicle, and a much smaller propellant weight, of a different type, with VASIMR. The unmanned vehicles would look the same. This path is a recommended possibility, although there is less commonality with the unmanned vehicles.

Opting instead for nuclear pulse propulsion runs into the odd efficiency scaling that kind of propulsion entails: Isp is higher at larger vehicle masses. At the masses of these exploration vehicles (619 metric tons at departure from LEO for the manned vehicle, 690 tons for each of 3 unmanned vehicles), pulse propulsion Isp resembles no more than gas core NTR Isp, and is maybe not as good. At vehicle masses around 10,000 tons and up, pulse propulsion looks like Isp = 10,000 sec or higher, and this at vehicle accelerations in the 2-4 gee range. These kinds of characteristics are well suited to large scale operations like base-building and planting actual colonies.

A Note on Crew Selection:

This mission is planned around 6 crew members, going to the surface 3 at a time. Since the goal is a science information return, to be obtained with maximum crew safety, I suggest each group of 3 be one pilot/engineer, one geology specialist, and one chemistry/biochemistry specialist. That would be two of each comprising the 6 total. Each should be cross-trained enough to function as a lander pilot for emergencies. Each should be cross-trained enough to support the other science specialties.

Concluding Remarks:

The point of this study was to show that a manned mission to Mars is feasible, and safe, with launch rockets available today or within 5 years.

All of the known crew health and survival issues can be addressed in a design that can be assembled from docked modules that fit the presumed launch rockets. Mission times are short enough not to provide artificial gravity. To go further out than Mars will require artificial gravity.

There are two missing propulsion technologies: (1) an update of the old solid-core NTR “NERVA” technology, which could probably be available in under 5 years, and (2) a gas core NTR (or VASIMR equivalent), which will likely require about 10 years to make ready.

These vehicle designs that support a well-planned exploration of Mars are reusable, and could be utilized anywhere in the inner solar system.

(see also the 1-8-2011 post for an update to this study)

Monday, November 29, 2010

Fast Transit To and From Mars

Based on some earlier work, I thought I would take a more systematic look at fast trips to and from Mars. I limited this study to gas core nuclear thermal rocket propulsion.

Orbital Tradeoffs for Transit:

The “typical” close approach distance from Earth to Mars at opposition is some 97 million kilometers. For very fast one-way flight times (on the order of 30 to75 days), the trip is very nearly a straight-line “shot”, at a rather low angle to a radial line. For fairly short stays at Mars (1 to 3 months), this picture doesn’t change much. Accordingly, I chose a nice round figure 100 million km, with an angle factor of 1. I looked at transit times from 30 to 75 days, computing average velocity as path length divided by transit time. A spreadsheet program automated these calculations. See fig.1.

Figure 1 – The Fast Transit Orbital Picture

The peak velocity is related to actual vehicle delta-vee. For fairly short impulsive delta-vee events, the velocity versus path length trace is nearly a square wave. The higher the acceleration, the closer the peak velocity is to the average. I assumed for this study the fairly brisk acceleration level of 0.5 gee at departure weight, knowing this would increase by the mass ratio at burnout weight, unless thrust were throttled back by the mass ratio. This takes the form of a manual convergence in the spreadsheet. For a round trip, vehicle delta-vee capability is 4 times this peak velocity.

Assumed Vehicle Characteristics:

The vehicle assumed for this generic trade study is depicted in the fig. 2 cartoon. This vehicle is single stage, 10% payload, 20% inerts, and 70% propellant. This corresponds to a vehicle mass ratio of 3-1/3.

Payload includes the crew, their habitat and supplies, their crew return vehicles, the Mars landing craft plus their propellants and all the exploration gear. All of this landing and exploration equipment is assumed to be returned to Earth, which is not realistic, but is very conservative. That “covers” the lack of a propellant “margin” in these figures.

Inerts include the vehicle connective structures, auxiliary electric power generating equipment, tank structures and insulation, cryo-cooler equipment to keep the liquid hydrogen liquid for months at a time, the basic gas core fission engine, and the waste heat radiator equipment which allows the engine to operate at Isp greater than about 2000 sec. In particular, the radiator could be a very major portion of that 20% inerts figure. It is unknown whether this is really adequate.

The propellant figure is the liquid hydrogen working fluid for the gas core fission nuclear thermal rocket engine.

Figure 2 – Basic Vehicle Assumptions

Velocity Requirements:

From the velocity trace modeled in the spreadsheet, we calculate average velocity versus one-way travel time, and peak velocity from the average using a constant 0.5 gee acceleration level. These are plotted in fig.3. At this level of acceleration, average and peak velocities are virtually indistinguishable. Required vehicle delta-vee for a one-way trip is twice the peak velocity, four times peak for a two-way trip, as plotted in fig. 4

Figure 3 – Average and Peak Transit Velocities

Figure 4 – Estimated Vehicle Delta-Vee Requirements

Assumed Propulsion Characteristics:

The focus here is on one type of propulsion: open-cycle gas core nuclear thermal rocketry. This is something that came very close to propulsive test in the late 1960’s.

At the end of the 60’s, controlled gas-phase fission had been demonstrated, as well as a powered-plasma flow model simulating a hydrogen to uranium flow ratio of 1000:1, which was as good as perfect containment. It was known that a waste heat radiator would be required, but not how it could be built, nor how much it would really weigh.

Neither the characteristics of that radiator, nor the maximum impulse capability of such systems, were known at that time. However, a first generation design concept was targeted at specific impulse Isp = 6000 seconds, because they were pretty sure they could achieve it. That is the screening criterion used in this study.

Estimated Propulsive Requirements:

From the mass ratio and the two-way trip delta-vee figures, the exhaust velocity required of the engine can be estimated. Dividing this by the gravity constant is a rather good estimate of the engine Isp required.

From the previous study, the payload for a crew of 6,twolanders,and two crew return capsules was 240 metric tons. Using this figure and the 10% payload fraction provides a constant estimate of Earth orbit departure weight, shown in fig.5 with the Isp.

Figure 5 – Required Isp and Departure Weight

The acceleration gee level at departure was assumed to be 0.5 for this study. At burnout weight on return, the acceleration is higher by the mass ratio, assuming constant thrust level. Under the constant thrust level assumption, the departure weight and acceleration set the engine thrust level. For these figures, the departure weight is 2400 metric tons, so the thrust level is 1200 metric “tons-force”, or 1,200,000 “kg-force”, which is some 11.76 mega-Newtons of force in proper SI units.

Thrust divided by Isp is propellant flow rate. Propellant mass divided by flow rate is the total burn time available to “dry tanks”. This was converted to hours for easy plotting. Vehicle departure and burnout gees, plus total burn time to dry tanks, is given in fig. 6.

Figure 6 – Acceleration and Burn Time Data

This oversimplified study is for a single-stage two-way trip: everything goes and returns, no expendables. From the standpoint of Isp’s thought feasible, it looks like such a vehicle could be powered by a gas core fission rocket, if the one-way travel time were no shorter than about 65 days. This is for an “average” opposition mission.

Possible major refinements: (1) reduce lander mass in favor of more lander propellants (extra landings) by taking some advantage of descent aerobraking, and (2) leave landers, unused lander propellants, and some small portion of the vehicle tankage behind in Mars orbit, as a depot for subsequent missions, thereby reducing payload and inerts for the return home.

Saturday, November 27, 2010

Mars in 39 Days One-Way !

This was a discussion topic in the forums on I decided to try a bounding calculation to rough-out what might be involved. My first attempt was not very good, but I have double-checked these numbers, They are good.

Orbits, Distance Traveled, and Times:

Mars and Earth move into opposition roughly every two years. The straight-line distance from Earth to Mars at opposition varies considerably, but averages 60 million miles (97 million kilometers).

For trips faster than about 60 days one-way, and stay times at Mars under about a month, the planets don’t have time to change relative position very much. Such travel is essentially a straight-line shot over a distance of about 60 million miles (see fig. 1). Longer travel times are different, because the orbits become curved, and the path along the orbit is much longer.

Figure 1 – Straight-Line “Shot” for Fast Trips Around Opposition

Constant-Acceleration Kinematics:

Along a straight-line path, at constant acceleration, the average and peak velocities can be estimated quite easily. Average velocity is merely the path length divided by the travel time. This is just high school physics math, but nothing better is needed.

I looked at two bounding cases: (1) vehicle acceleration so low that one thrusts to accelerate to the mid-point of the one-way path, skew-flips under power, and then thrusts to decelerate the second half, and (2) one accelerates “hard” at each end such that time spent in acceleration is only 10% of travel time for each of the two “burns” (the “10% tails” case).

Both cases put lower limits on vehicle thrust. The continuous-burn skew-flip case has a lower ratio of average velocity to peak velocity at 50%. The 10% tails case has average velocity 90% of the peak velocity. See fig. 2.

In both cases, one computes the acceleration required as the peak velocity divided by the burn time. The two peak velocities are different for the two cases. The skew-flip burn time is half the travel time. The “10% tail” burn time is 10% of the travel time.

Figure 2 – Kinematics for Straight-Line “Shot” Trips: Two Cases

The data inputs here are path length 97 million km, and 39 days travel time. This figures to an average velocity of 28.79 km/s. To go there and back again, requires a total of 4 burns, each to the peak velocity.

My figures from the hand estimates and from the spreadsheet are a peak velocity of 57.57 km/s for the skew-flip option, and 31.99 km/s for the “10% tails” case. These correspond to total propulsive delta-vee requirements of 230.3 km/s for the skew-flip option, and 127.9 km/s for the “10% tails” option.

The 10% tails option spends most of its time coasting at peak velocity, reflected in the higher average/peak velocity ratio. For the same average velocity, the “10% tails” option has a lower peak velocity, and a lower delta-vee requirement, by a substantial amount. However, the thrust levels demanded are substantially lower for the skew-flip option, since it has half the trip time to accelerate, instead of 10% of it.

Vehicle Components and Configurations:

In order to do a little better than a simple Apollo-style “stunt” landing, I decided to send a fairly substantial crew of 6 with two landing vehicles and two crew return capsules. I included a habitat module on the order of the old “Skylab” module. The idea would be to rotate the crewmembers on successive surface expeditions, 3 in each lander, to cover 2 separate landing sites.

My estimates for the habitat reflect stored supplies, no artificial gravity, but a radiation-sheltered room from which the ship may be maneuvered during a solar flare event. I simply guessed this module at 90 tons, same as the old “Skylab” station. There is enough volume in a module this big for a crew of 6 to stay sane and healthy for perhaps 6 months at a stretch. The design here is 3-4 months round trip.

The crew return vehicles would each seat 3 comfortably, all 6 in a pinch, rather like the original Apollo capsules. These would have an abbreviated service module with hypergolics propulsion for deorbit maneuvers. I simply guessed these at 5 tons each, for 10 tons total. See fig. 3.

Figure 3 – Payload for the Orbit-to-Orbit Transfer Vehicle

I wanted to rough-out the landers as single-stage nuclear devices. This would provide the option to leave these vehicles in orbit at Mars, to be used again.

Mars surface-to-orbit requires a little under 6 km/s delta-vee one-way. Accordingly, I roughed out a single-stage rocket for a total 12 km/sec delta-vee, using a solid core nuclear thermal Isp = 1000 sec. This allowed me 9% payload, 71% propellant, and a 20% inert fraction (to cover landing legs, a cargo crane, and maybe a bit of shielding around the reactor.

I figured the payload as 3 men and supplies plus some sort of a rover car, similar to what we did on the moon. I guessed 6 tons for this, and calculated just under 70 tons for each lander. For two landers, that adds 140 tons to the transfer ship payload. See fig. 4.

The old NERVA solid core engine tested as 1000 sec, with an engine thrust/weight near 3. This was insufficient for surface launch from Earth, but might work on Mars. A design update to incorporate the results from projects like Dumbo and Timberwind would very likely provide a 1000 sec Isp engine with thrust/weight 10+.

Figure 4 – Single-Stage Solid Core Nuclear Thermal Lander

I looked at a variety of configurations, based around a core vehicle with a single engine or engine cluster, and some number of added propellant tanks (with no engines of their own). The idea would be to assemble this vehicle in low Earth orbit (LEO), and fly it to Mars orbit, then back to Earth orbit. Maximum recovered components provide maximum reusability for future missions.

I looked at a very conservative design for the core vehicle, to cover both the robustness necessary for reusability, and the connective structure to hold propulsion, tanks, and payload items together. I assumed 10% payload fraction (the 240 tons discussed above), and 20% structural inerts fraction. That leaves 70% for propellant load.

The tanks are also conservative, in that I assumed 10% structural inerts and 90% propellant. This should cover some sort of cryo-cooler equipment items necessary to store liquid hydrogen as propellant for the presumed nuclear engines. See fig. 5.

Figure 5 – Assumptions for Core Vehicle and Added Tanks

Not every configuration I looked at uses staging (in the sense of dropping empty tanks). For those that did, I assumed only two burn phases: (1) burn all the propellant out of all the tanks and then drop the empties all at once, and (2) burn on the propellant in the core vehicle. This is suboptimal, as one could drop tanks as they become empty (it might be necessary to do this in pairs to preserve weight and balance). Accordingly, my estimates of assembled vehicle weight and propulsive Isp may be a little bit high, but only by a very few percent. See fig. 6.

The list of configurations is as follows:

(1)baseline core vehicle without tanks, as a single stage (all recovered and reusable)

(2)core vehicle plus tanks at 50-50 weight split, single stage (recover and re-use tanks as well as core)

(3)core vehicle plus tanks at 50 - 50 weight split, two stage (expendable tanks)

(4)core vehicle plus tanks at 20 core - 80 tanks weight split, two stage (expendable tanks)

(5)core vehicle plus tanks at 5 core – 95 tanks weight split, two stage (expendable tanks)

Figure 6 – Staging Assumptions for Expendable-Tank Configurations

All of these configurations use the same core vehicle and 240 ton payload assembly. The transfer propulsion is on the core vehicle. It is assumed to make all four burns required by each round trip, and be reusable for several more trips.

I ran a weight statement for each of these configurations, based on the component weight fractions, the shipped payload (240 tons), and the core-tank weight splits.

For each of the three staged configurations, the mass ratio for the first delta-vee burn is departure weight divided by departure weight minus the total tank propellant weight. The mass ratio for the core vehicle is always the same: core vehicle weight divided by core vehicle weight minus core vehicle propellant weight. Assuming the same Isp for both burns (it is the same engine), one may sum the natural logarithms of these mass ratios and use that sum as total vehicle delta-vee divided by engine exhaust velocity.

The two single stage configurations are the core vehicle alone and the core vehicle plus tanks that are never dropped. One simply runs single mass ratios based on the weight statements for these configurations.

These mass ratio-based velocity ratios may be used with the total delta-vee requirements from the kinematics to estimate propulsive Isp. Delta-vee divided by the velocity ratio is an estimate of engine exhaust velocity required to perform the mission. This exhaust velocity divided by the gravity constant gc (9.805 m/s2 = 32.174 ft/sec2) is a rather good estimate of required propulsive Isp.

From the 240 ton payload weight, and each configuration’s weight statement, one can estimate the departure weight of the vehicle that must be assembled and fueled in LEO. From that departure weight and the average acceleration required by the assumed kinematics (skew-flip or “10% tails”), one can estimate the minimum allowable actual thrust required of the engine system.

Those data are given in fig. 7. Thrust figures are minimum allowables. Remember, for the skew-flip option, a minimum vehicle acceleration of 0.0035 gees is required; higher values allow one to coast a little between the acceleration and deceleration burns. For the “10 tails” option, a minimum vehicle acceleration of 0.0097 gees is required; higher values simply shorten the “tails”.

Figure 7 – Results for 5 Configurations, and for 2 Kinematic Options

Indications of Propulsion Required:

The single-stage core-only configuration has a relatively low assembled weight in LEO (although still far too high for a single launch), but requires propulsive Isp beyond anything we understand with nuclear thermal rocketry, whether one looks at the skew-flip or “10% tails” scenario.

A variable-thrust electric approach such as VASIMR might possibly serve, although there are very serious questions whether an 8-24 ton thrust VASIMR can be powered by something that would fit this weight statement. For all the skew-flip configurations, a higher thrust phase is required for efficient orbital escape and capture at each end of the trip, something easily handled by VASIMR.

The single stage core plus non-expendable tanks, at a 50-50 core-tank weight split, has the effect of improved mass ratio and reduced impulse requirements. The 4800 ton departure weight is something that seems “reasonable” for assembly and fueling in LEO, although it is clear that many launches would be required.

For both the skew-flip and “10% tail” scenarios, the impulse requirements seem infeasible even for gas core nuclear thermal rockets (maximum in the 6000-10,000 sec range). The higher thrust levels make the VASIMR power supply problem worse, even in the heavier vehicle.

The two-stage core plus expendable tanks, at a 50-50 core-tank weight split, has the same departure weight of 4800 tons. The impulse requirements for the skew-flip and “10% tails” scenarios are reduced a little, but only a little bit. Even the “10% tails” impulse requirement is likely still infeasible for a practical first-generation gas core nuclear thermal rocket. A waste heat radiator is required, and would be rather heavy. It might not fit well in this weight statement.

The two-stage core plus expendable tanks, at a 20-80 core-tank weight split, has a larger departure weight of 12,000 tons. This is getting very ambitious for assembly and fueling in LEO. However, for the 10% tails scenario, the thrust level and specific impulse requirement looks very achievable for gas core nuclear thermal rocketry. The skew-flip scenario’s requirements still look infeasible for this type of propulsion.

The two stage core plus expendable tanks, at a 5-95 core-tank weight split, provides only a small further reduction in impulse, at the cost of an enormous increase in departure weight. 48,000 tons assembled and fueled in LEO is a rather daunting prospect, for only a reduction from 5270 sec to 4160 sec gas core Isp in the 10% tails” scenario (skew flip is still infeasible). There is no option here for solid core nuclear thermal main propulsion, and little prospect for it by extremizing the weight statement still further.

I would hazard a guess that solid core nuclear thermal propulsion would only become feasible if we multi-staged the core vehicle, including the nuclear engines. This is not an attractive possibility.

On the other hand, nuclear pulse propulsion is a method that is feasible, just not well-matched to payload-driven weight statements like these. “Small” systems in the 2000-5000 ton departure weight regime would have Isp values in the 2000-5000 sec range, based on the old design data from about 1960. Much larger systems, in the 10,000 to 20,000 ton range would have Isp values in the 10,000 to 20,000 sec range.

A large pulse propulsion vehicle in the 10,000 to 20,000 ton departure weight range would have the impulse to achieve the 39 day mission. The 2-4 gee vehicle accelerations available would cut the “10% tails” to 1% or less. Deliverable payloads would be more in line with colonization efforts that initial explorations, though, being around 10 or more times the 240 ton figure used here.


The “10% tails” scenario is preferred because of the substantially lower impulse requirements, in spite of its higher minimum thrust requirements.

The impulse levels reported here are being driven by the very high average velocities associated with a 39 day one-way timeline for the flight. These vehicle results are being driven very hard by those impulse figures.

If the electric power plant problem can be solved, then a 4800-or-larger ton VASIMR-powered vehicle operating on the “10% tails” scenario might be an attractive option. This would be a staged vehicle with expendable tanks, at a 50-50 weight split. The VASIMR would have to operate briefly at lower Isp and higher acceleration to effectively “get away”, then throttle back to higher Isp at low thrust, until the required peak velocity is attained. That same variable-performance scenario in reverse is required for effective capture into Mars orbit. A similar profile is required for the return.

If the high-temperature radiator problem can be solved, then a 12,000-or-less ton vehicle powered by a gas core nuclear thermal rocket operating on the “10% tails” scenario can do this mission. The impulse is slightly lower than the 6000 second design target of the ca.-1970 design data, and the thrust is in a feasible range.

Other Options to Consider:

Mass ratio and delta-vee can be improved slightly with any of these designs by leaving the lander craft behind in Mars orbit for future re-use. Subsequent missions then need only bring propellants for these landers, not the entire craft, up to the design lifetimes of the landers’ engines.

Leaving the rover cars and other exploration gear behind would ease lander propellant consumption slightly, and provide ready assets for a subsequent return to these sites.

It might be wise to re-run this analysis at a 60 day one-way time, in order to see just how strongly that parameter drives these vehicle designs.

It would be wise to explore just how best to design a pulse propulsion vehicle to do this mission. Such designs are not driven by payload weight (mass ratio-limited at a separately-defined Isp), but by departure weight (which actually defines the Isp).

Detailed Weight Statement Data Used:

1-stage core only
Payload......0.10.......MR = 1.00/(1.00 – 0.70) = 3.333333333
Propellant...0.70.......ΔV/Vex = ln (MR) = 1.20397280

Core + non-expendable tanks (50-50)
Payload......0.10–0.05..0.00–0.00...0.05.....MR = 1.00/(1.00 – 0.80) = 5

Core + expendable tanks (50-50)
Component....core.(50%).tanks.(50%).overall...MR1 = 1.00/(1.00 – 0.45)
Payload......0.10–0.05..0.00–0.00...0.05..........= 1.818181818
Inerts.......0.20–0.10..0.10–0.05...0.15......MR2 = 0.50/(0.50 – 0.35)
Propellant...0.70–0.35..0.90–0.45...0.80..........= 3.333333333
ΔV1/Vex = ln (MR1) = 0.597837001........ΔV2/Vex = ln (MR2) = 1.20397280

Core + expendable tanks (20-80)
Component....core.(20%).tanks.(80%).overall...MR1 = 1.00/(1.00 – 0.72)
Payload......0.10–0.02..0.00–0.00...0.02..........= 3.5714
Inerts.......0.20–0.04..0.10–0.08...0.12......MR2 = 0.20/(0.20 – 0.14)
Propellant...0.70–0.14..0.90–0.72...0.86..........= 3.333333
ΔV1/Vex = ln (MR1) = 1.2729657..........ΔV2/Vex = ln (MR2) = 1.20397280

Core + expendable tanks (5-95)
Component....core.(5%)...tanks.(95%).overall...MR1 = 1.000/(1.000 – 0.855)
Payload......0.10–0.005..0.00–0.000..0.005.........= 6.89655
Inerts.......0.20–0.010..0.10–0.095..0.105.....MR2 = 0.050/(0.050 – 0.035)
Propellant...0.70–0.035..0.90–0.855..0.890.........= 3.333333
ΔV1/Vex = ln (MR1) = 1.93102115...........ΔV2/Vex = ln (MR2) = 1.20397280

Wednesday, November 17, 2010

Nissan Mileage Results on Blends

These data came from over a year's worth of stiff blend trials in the unmodified Nissan. This was a 1998 "Sentra". I had no way to draw fuel samples until the last two tanks of fuel, so the blend strength data are best-estimate E-numbers for what was added at each fill-up. All but one or two of these were fill-ups on top of quarter tank-or-less residuals, so it's a decent estimate, no matter what.

Although scattered at about plus-or-minus 10%, there does seem to be a very slight (5%) decline in fuel mileage between all gasoline and 25-something-% blends. Maybe. The decline is 5% when the scatter is plus-or-minus 10%. Maybe that drop is not actually real.

Theoretical lower heating value for an E-25 is 91% that of plain E-0 gasoline: a 9% drop. It would seem the car is doing better than heating values would indicate, although the scatter is too great to actually draw that conclusion. This is pretty much the same result as was seen with blend trials in the unmodified Ford F-150.

One tank (only) was accidentally too rich in ethanol at an estimated E-49. This one data point is about 80% of the E-0 mileage. Such a drop is big enough to be "real", even with plus-or-minus 10% scatter. There are points without drop at about E-40 to E-42. This behavior is consistent with stiff blend trials in the unmodified Ford, and with the ethanol VW as rigged to gasoline settings. (See the November 12, 2010 post just prior to this one). The VW data are spot-check "snapshots" from intake vacuum measurements, not tank averages of mileage.

That's three wildly-different cars providing the same quantitative answer, and from two different completely-different types of data. Stiff blends in unmodified vehicles provide essentially gasoline-only performance up to blend strengths in the E-40-to-42 range.

For the Nissan, I was able to cross-check the best-estimated E-numbers against actual test measurements on the last two tanks. Both were within two percent of being correct.

I see nothing here that contradicts the conclusions in the earlier post:

(1) Above E-42, timing seems to be late (low vacuum, low performance).

(2) Below E-42, you cannot practically tell these blends from a straight gasoline.

Considering that no sophisticated test equipment was used at all, these results are quite remarkable, are they not?

Friday, November 12, 2010

Stiff Blend Effects in Gasoline Cars

I recently completed the final "stiff blend" experiments with the "ethanol VW". This was the 1973 VW Type 1 sedan that I had converted to E-85-only operation in January 2007. In fall 2010, I had re-converted this vehicle to gasoline-only settings, and tried a series of increasingly-stiff (high E-number) blends. To make such testing practical, I added an adjusting screw to the E-85 main jet to make this a "flex-fuel" carburetor. This old-technology vehicle is carbureted, with distributor ignition.

For these stiff blend experiments, I used no added intake air preheat, and no extra timing advance. For most of these runs, I used the gasoline-only jet settings on the main and idle jet screws. I used the E-85 accelerator pump discharge nozzle as it was. Later in the tests, at higher E-number blends, I reset the main and idle jet screws a little, to accomplish manually the same "compensation" that electronic fuel injection does automatically. This was necessary to maintain good drive-ability.

I tested this VW in terms of intake vacuum observations during in-gear coast-down. This was based upon the well-known fact that late ignition timing shows up as lower intake vacuum levels, all else being equal. Blend strength was determined by a simple forced added-water phase separation test. The theory behind this is that ethanol requires more advanced ignition timing, because it has a longer ignition delay. Thus, you have to "light the fire" sooner, in order to achieve peak in-cylinder pressures by about 2 degrees after top dead center crankshaft position.

What I found was absolutely no differences between gasoline-only and stiff blend performance up to about 40% ethanol in the mix (E-40 blend). For blends from there up to the maximum tested (E-56), there was a definite drop in coast-down vacuum, although it was not large. My data curves show a definitely-repeatable "shape" that is most likely a systematic gage-reading error in these on-the-road tests. Essentially, all these blends show just about the same vacuum loss, once 40% was exceeded.

The VW blend experiment was inspired by earlier results in 2008 with stiff blends in my 1995 Ford F-150 pickup truck. This vehicle has distributor ignition and electronic fuel injection. I ran these stiff blends in it completely unmodified, and I still do today. The data for the truck were obtained from fuel mileage, not intake vacuum. I kept track of two driving cycles: "TSTC", which was a virtually-all highway commute 30 miles one-way to / from work at TSTC-Waco, and "SRMcL", which was general driving over shorter trip ranges around McLennan county, Texas.

Driving cycle on blend fuels did not seem to make much difference, after normalizing blend mileage to the corresponding gasoline-only average mileages on the same driving cycles. Somewhere just above E-40 blend strength, fuel mileage dropped dramatically. The performance characteristics above 40% ethanol were very smooth operation, low perceived power, and the sense that it was "sucking fuel too fast", confirmed later by the actual mileage data.

The data shown here are exactly the same data that were in the February 2008 "Ford Report 8" (see my website, and navigate to the "ethanol projects" sub-page). A few more data points added later (by May 2008, but not included here) confirmed the same basic trend, and "pinned down" the critical blend strength a little better, to right at E-42.

The most amazing thing about this is that two vehicles of such different vintages and technologies gave almost exactly the same answer from two completely-independent types of data. Timing is late on gasoline settings, in otherwise unmodified vehicles, above about E-40-to-42 blend strength.

I got pretty much the same answer, although not actually quantified, from a 1998 Nissan Sentra early this year. My best estimate was that I accidentally put a blend a little above E-50 in the tank, causing perceived low power, low mileage, but smooth operation (exactly what one would expect from late timing).

It looks to me as if above about E-40 to E-42 blend strength in an unmodified car, the timing is suddenly late, as if "turned off like a light switch". Up to that point, one cannot in any way tell stiff blends from straight gasoline (either by mileage or by intake vacuum).

I would speculate that in blends under the critical level, the gasoline component controls the effective ignition delay, while for blends over the critical level, the ethanol component controls the effective ignition delay. Thus for stiff blends, the required timing setting is not smoothly dependent upon blend strength, but instead is a step function triggered by a critical blend level near 40-42% ethanol.

This finding has serious implications for the design of factory flex-fuel vehicles.

You heard it here, from me, first!

Monday, October 25, 2010

Election 2010: TX Governor

My 3 issues for choosing are:

1. Factual sources of information (most definitely not campaign ads).
2. Re-elect no incumbent that you cannot personally verify did more good than harm.
3. Look for real specifics in proposed actions(names and numbers, not slogans).

Factual data as reported by the newspapers:

Governor Perry was behind the attempted land grab for the Trans-Texas Corridor, a proposed foreign-owned toll road. He vetoed the eminent domain reform bill passed by the Legislature specifically to make this possible. At stake was almost a million acres of privately-owned land to be forcibly taken by the state, and turned over to a foreign company for their profit. The bill he vetoed specifically outlawed this taking of private lands for private profit.

He was also behind the attempted fast-tracking of permits for a big cluster of old-technology coal-fired power plants to be built close to the greater Waco area. This would have driven McLennan county into air pollution non-attainment status, forcing local residents to get their cars emissions-tested for state inspection, at increased expense. It would have also driven the Dallas-Ft. Worth and Houston non-attainment areas into direct federal EPA control, something no one here wants.

More good than harm:

The two items listed above would have done great harm to the people of Texas. I see no way that the good Mr. Perry has done, and he has done some, outweighs these two items. He fails, and Mr. White wins, on this selection issue.

Proposed specifics:

I have not yet seen much (that I trust as factual data) from Mr. Perry or Mr. White in the way of specific policy proposals. It's a "wash".


I recommend voting for Bill White for governor. That's what I will be doing.

Saturday, October 23, 2010

On Election 2010

It’s election time, which is always a very important decision, one I encourage you to make. Allow me to share three ideas, which I heartily recommend you consider, as you make your decisions.

First: believe only sources of factual information, campaign ads and internet forwards most definitely do not qualify. Stick to public debates refereed by responsible parties, and to interviews by actual professional journalists, as reported in the newspapers.

Do not place much weight on simple speeches. Those are good for drawing your attention, not so much for honest, factual, and substantive information.

Second: there is the tendency of incumbent politicians to become corrupted as time-in-office passes, in the sense that special interests buy them with lots of financial contributions. This takes time to occur, so as a general rule, regular turnover insures better, more honest representation for the common man.

There are exceptions to this generalization, but only down at the 1-3% level. If you can verify for yourself, from factual information, that an incumbent did more good than harm, then he is worth re-electing.

Third: when comparing two candidates, look at the details of what they say, not just the “sound” of it (which is really your own ideological “filter” talking, rather than anything the candidate actually said). Look for names, numbers, and other specifics. If you do not see specific proposals with actual figures and names, then you are seeing only slogans and ideology.

Better to choose the candidate with substantive ideas than the sloganeer. Ideologies and slogans historically made very bad public policy, such as happened in the former communist world, among many other examples.

In my book, looking at candidates with those three issues in mind trumps any possible politics, be they the candidate’s or yours. What you are looking for is an honest, thoughtful person, with substantive ideas, who will do right by you.

You choose for yourself. But I do offer these two thoughts:

I recommend re-electing Chet Edwards. He has by far the more substantive ideas about what to actually do. And by voting against his own party at times, I really do know Chet does more good than harm.

I also recommend re-electing Doc Anderson, and for the same reasons. Doc voting against his own party is partly why our kids will not grow up under a pall of coal ash. That is most definitely more good than harm.

As for the rest? I haven’t made up my mind yet. But, if I don’t know, then I vote “no”. A lot of them have not made their cases to me.

Friday, September 10, 2010

More Troops and Fences on the Border Won’t Fix Our Troubles

I get all kinds of obvious political “hit pieces” forwarded to me by lots of friends and neighbors. These deal with all kinds of things; one of the more frequent subjects is illegal immigration from Mexico. What follows is a typical forward of a video on YouTube that implies all our border violence and danger troubles are due to illegal immigration. The facts in the piece are true, it is what gets left out that is so very telling. I tried to debunk the “hit piece” by supplying as much of the whole truth as I understand, in my response.

From a friend in Austin, dated 9-1-10:
This is a crying shame! When is Washington going to get off its ass and do something? Making them all citizens is stupid! Are we not going to enforce the drug laws either? I heard the other day the they are complaining that Austin and Travis County is the biggest location for deporting "undocumented aliens" who haven't broken any laws in the country! Excuse me, isn't being an "illegal alien" a crime??

The link was to a video of a Texas Congressman blaming all the border violence on illegal immigrants and drug cartels from Mexico. He got it half right: the cartels are indeed causing a public danger to Americans. But not so much the immigrants. I do agree with his criticism of the federal government for shirking its duties in this area. I disagree that more guards and more fences would stop either the drug runners or the illegal immigrants.

This was my response to my friend, dated 9-2-10:

What you're talking about and what's on the video is all true, but not quite complete. The drug violence that is now spilling across the border is fueled primarily by demand for illegal substances here in the US. There is a drug "industry" in Mexico that thrives only because it can command high prices for its illegal product. Those prices are high precisely and only because those substances are illegal (we've seen that effect before, with alcohol during Prohibition). And that's their Achilles heel, too.

Another piece of that puzzle is that there have been drug cartels in Mexico plying this trade for many decades. There is a history of ever-more violent competition among those cartels all during those years. It has really extremized by the current time. Ever since about 1970, the cartels were better armed than the Mexican military. I'm talking armed jet aircraft, tanks, machine guns, and heavy weapons, all funded by drug revenues. Most of that is supplied by no-questions-asked arms dealers on the US side of the border. Our border guards never looked (and still do not look) for weapons being taken into Mexico. Stopping the arms supplied to the cartels is what Calderon wanted to talk about while he was up here a few months ago. It's still a wide open flow of heavy weaponry south, and they still outgun the Mexican military today, by a longer shot now than ever before. Until that flow of arms is stopped, the Mexicans cannot win their drug war, and the violence spilling across the border cannot be stopped, on either side.

Now, did you notice that absolutely none of my discussion has anything to do with immigration, legal or not? Any illegal immigrants who commit violent acts (and there are some) are very tiny small change compared to the drug war problem in Mexico. Stopping entirely the flow of illegal immigrants (or even all immigrants whatsoever) will not stop the violence spilling over, and has nothing to do with the cartels that are the real problem. There is no border fence that cannot be breached. Blaming all this violence on illegal immigration is a political campaign fiction, pure and simple. It has nothing whatsoever to do with the truth of the matter, or with any reasonable plans we might make in order to deal with it.

What needs to be done about the spillover drug violence is two-fold: (1) stop the southward flow of weaponry to the cartels (which is clearly a higher priority than stopping any northward flow of immigrants, legal or not), and (2) kill the revenue streams of the cartels by legalizing, regulating, and taxing the drugs they sell. Made legal, prices fall, quite dramatically. If the cartels can't make scads of money at it, then they get out of the business, and that problem goes away. It won't even take very long: the Mafia was out of bootleg booze within a year or so of Prohibition's end, and all the bootlegging violence died with that change.

The illegal immigrant problem is pretty much separate from, and insignificant compared to, the cartel drug violence problem. No amount of campaign sloganeering, posturing on the floor of Congress (the video you forwarded), or political party propaganda is going to change that simple truth. But facing and telling the truth does not win votes at this time in this country, scapegoating groups that cannot defend themselves does, just as it did in 1930's Nazi Germany! And that's exactly what the right-wing extremists in this country are doing, and for several years now.

I quite agree that illegal immigrants shouldn't be here. I quite agree that we need to get the illegals out of this country. I disagree with a lot of schemes to do that, but there is a workable and fair way, and it will be found. I disagree that illegal immigrants are the cause of "all our troubles" (really meaning border violence). I disagree that securing our borders against illegral immigration is a "top priority". Yes, it needs to be done, but the top priority is ending the drug cartel war in Mexico, and I already told you how to do that, and how it has almost nothing at all to do with illegal immigration.

It is seeing that scapegoating thing going on right now, in this country, that scares the bejeezus out of me. History is repeating itself, and that is one ugly history.