In Part 2 of Andrew Higgins’ discussion of laser-thermal rocketry and fast missions to Mars, we look more deeply at the design and consider its potential for other high delta-V missions. Are we looking at a concept that could help us build the needed infrastructure to one day support expansion beyond the Solar System?
by Andrew Higgins
We now turn to the detailed design our team at McGill University came up with for a laser-thermal mission capable of reaching Mars in 45 days. Our team took the transit time and payload requirement (1 ton) from a NASA announcement of opportunity that appeared in 2018 that was seeking “Revolutionary Propulsion for Rapid Deep Space Transit”. Although being in Canada made us ineligible to apply to this program, we adopted this mission targeted by the NASA announcement for our design study; being in Canada also means we are used to working without funding.
Image: McGill University students responsible for the design of the laser-thermal mission to Mars.
The NASA-defined payload of 1 ton would be a technology demonstration mission (what we call Mission Mars 1 in our study). Placing a premium on minimizing the transit time presumably reflects NASA’s eventual interest in lessening astronaut exposure to galactic cosmic rays, which increases sharply once a spacecraft leaves the Earth’s protective magnetosphere. Once on the surface of Mars, data from the Curiosity rover have shown that the radiation environment there appears to be more benign, comparable to or even less than the radiation exposure encountered on the ISS. Throwing regolith to cover the habitat on Mars would lower the radiation risk further, so astronauts leading a hobbit-like existence on Mars should stay healthy, provided they get there quickly.
Our Mars 1 mission starts with our spacecraft already in medium Earth orbit (MEO), so that it remains in view of the ground-based laser during the entire laser-powered burn, which takes about an hour. Given the ongoing revolution in space access, we did not bother to explore using laser propulsion to get to orbit. Chemical propulsion is well-suited for reaching orbit, so we selected a Falcon 9 to bring our vehicle to MEO and focused on using the laser for the transit to Mars.
Image (click to enlarge): The concept of operations for a rapid transit to Mars mission using laser-thermal propulsion. Note the use of a burn-back maneuver to bring the laser-thermal stage back to medium Earth orbit after sending the payload to Mars.
The laser array on Earth is about 10 m by 10 m, comparable to a volleyball court, and for the 1 ton payload mission, the laser would operate at 100 MW output for an hour, using power taken from the grid or generated via solar and then stored in a battery farm. (It is worth noting that a battery farm capable of providing 100 MW for an hour was built in South Australia in 2017 from scratch in just 60 days, in response to a taunt posted in a tweet [1]. So, powering the laser is not a problem.
When the laser beam arrives at the spacecraft, it is focused into the propellant heating chamber by a large, inflatable reflector—a balloon that is transparent on one half and reflective on the other. Inflatable space structures like this are fairly mature, including a demonstration of an inflatable antenna that flew on the Space Shuttle in 1996; a comprehensive overview of this technology was given by Jamey Jacob at the 6th TVIW in Wichita [2]. Inflatable collectors such as these have shown sufficient optical quality for our purposes. While the laser flux on the inflatable is intense, we found fluorinated polyimide films have sufficiently low absorptivity to avoid overheating.
Image: Inflatable Antenna Experiment deployed from the Space Shuttle Endeavor (STS-77).
Image Source: https://apod.nasa.gov/apod/ap960525.html
The inflatable reflector focuses the laser into the heating chamber, raising the temperature of the hydrogen flowing through the chamber to greater than 10,000 K. Keeping the walls of the chamber cool is the central challenge of the design, but our team found a combination of regenerative cooling (cool hydrogen flowing through the walls), transpiration cooling (injecting hydrogen through porous walls), and seeding the hydrogen (to trap thermal radiation in the propellant, similar to the greenhouse effect) should be sufficient to keep the walls cool. The heat absorbed via regeneration is used to power the turbopumps needed to pump the hydrogen via an expander cycle. The fully ionized hydrogen propellant is then exhausted through a conventional bell nozzle to generate thrust. Based on our own calculations and prior work on laser thermal propulsion and gas-core NTRs from the 1970s, a specific impulse of 3000 s appears feasible.
Image: Details of the propellant heating chamber and associated propellant feed and cooling systems.
The laser propulsion hardware is just dead mass once the spacecraft exceeds the focal length of the laser (which is about 50,000 km), so our team proposed bringing the laser thermal propulsion stage back to Earth via a flip-and-burn-back maneuver while still within range of the laser in cis-Lunar space. Once the propulsion stage is brough back to low or medium Earth orbit, it can be refilled and readied for use again. This would allow a single laser-thermal stage to throw multiple payloads to Mars over the duration of a given launch window.
The 14 km/s Delta-V laser thermal burn sends the spacecraft to Mars on a nearly straight line trajectory: no need for looping ellipses and Venus flybys. Our astrodynamicist optimized the trajectory for a 2020 departure. Even though our design had the launch two months after Perseverance, the vehicle would arrive at Mars three months before the newest Mars rover, overtaking it on the way.
Image: 45-day transfer orbit to Mars via laser thermal propulsion, in comparison to the 7-month journey of the Perseverance rover.
When the spacecraft arrives at Mars, there is no laser to perform a laser-assisted deceleration burn (at least, not yet) and at the high approach velocity, aerocapture appears the best option. At an approach speed of 16 km/s, aerocapture is going to be harsh and is another critical link in the mission design. The heat flux will be intense, but the new Heatshield for Extreme Entry Environment Technology (HEEET) developed by NASA in recent years appears to be rated to withstand even greater heat flux. The vehicle entering the Martian atmosphere would need to use lift pointed down (toward the surface of Mars) to keep the vehicle in a trajectory that skims the atmosphere. This maneuver is a delicate balance between heat load, the g-load, and the lift and ballistic coefficients of the spacecraft, which we first modelled analytically and then backed-up with full three-degree-of-freedom simulations. The g-load limit was set at 8-gees for our study; for the scaled-up design with astronauts, the g-load will be severe and sustained for several minutes, but within the limits of what humans can tolerate. (Relevant to note that, at the recent Interstellar Symposium in Tucson, Esther Dyson reported from her centrifuge training at Star City that, “8-gees going through you was actually a lot of fun” [3]). The aerocapture would be a wild ride, for sure.
Image: Details of model used for aerocapture upon arrival at Mars.
The scaled-up version of our design (Mission Mars 2a) intended for crewed missions used a 40-ton spacecraft derived from the Orion capsule and European Service Module. The greater payload requires a more powerful (4 GW) laser to effectuate the same 45-day transit to Mars, but the laser array occupies the same 10-m footprint on earth.
The other mission we considered was a cargo mission (Mission Mars 2b). Robert Zubrin often makes the point that—even if advanced propulsion capable of high thrust and high specific impulse was available—he would still opt for a 6-month free-return trajectory and use the enhanced propulsion capability to bring more payload. So, the Mars 2b mission uses the performance of laser thermal propulsion to maximize the amount of cargo that could be brought to Mars with a Hohmann-like transfer, and shows that the payload could be increased by a factor of more than 10 over what a Centaur upper stage—with the same mass of propellant—could throw to Mars.
Image: Final design of laser-thermal propulsion spacecraft capable of reaching Mars in 45 days.
While a more thorough vetting of our design is called for and much work remains to be done, one encouraging finding is that the specific power of the laser thermal propulsion design is so good—an “alpha” on the order of 0.001 kg/kW—that even if the mass of the entire propulsion system were to increase by a factor of ten, the increased mass would not significantly affect the overall performance or payload capacity of the design. There is sufficient margin in the concept to accommodate the inevitable upward creep in mass that occurs as the design is refined.
Laser thermal propulsion may be well suited to other high Delta-V missions, such as flybys of interstellar comets, the mission to the solar gravitational focus, and a probe to the hypothetical Planet 9—if it is found. There is no reason the laser-thermal approach cannot be combined with laser electric propulsion or other techniques such as an Oberth maneuver. Perhaps it is best to think of laser thermal propulsion as a dragster that burns a lot of propellant quickly to get you up to speed, but from there, you can invoke laser electric propulsion that is well suited to the diminishing laser flux as the spacecraft exceeds the focal length of the laser. Appendix A in our paper details where we calculate the tradeoff between laser thermal and laser electric propulsion occurs. Hopefully, the laser-thermal concept can contribute to a further appreciation of directed energy as a disruptive technology for high-velocity missions in the solar system and beyond.
The complete details of our study can be found in our published paper: Duplay et al, “Design of a rapid transit to Mars mission using laser-thermal propulsion,” Acta Astronautica Volume 192 (March 2022), pp. 143-156 (abstract / preprint).
A browser-friendly version of the paper is available here: https://ar5iv.org/html/2201.00244
References
1. https://www.popularmechanics.com/science/a31350880/elon-musk-battery-farm/
2. J. D. Jacob, B. Loh, Inflatable technologies for interstellar missions, in: P. Gilster (Ed.), Proceedings of the 6th Tennessee Valley Interstellar Workshop, 2020.
3. https://www.youtube.com/watch?v=nHnUeM8RovE
I think more needs to be said about how the design handles the return burn to MEO. The craft will need to decelerate. In that orientation, the thrust chamber would be firing into the direction of travel to cancel out the delta_V. However, the mirror would now have its rear surface facing the beam and its positioning wrt to thrust chamber window would be incorrect.
If the thrust chamber window was on the “top” of the thrust chamber. the mirror would just need to rotate about its axis over the window to facilitate this. Yes, the mirror would need to be a little larger to intercept the full 10m beam, but it would make this maneuver much easier to design for. That this solution was not chosen makes me wonder if there is some constraint on the placement of the window on the thrust chamber that requires it to be placed where it is in the reference design.
Once a functioning Mars base is built, it seems that a laser array near the base would be useful so that the craft makes the full journey to Mars, is decelerated to orbit, and then refueled for a trip back to Earth. One now has a fast, 2-way service that might be needed for at least some craft to make the full return journey to return payloads from Mars. If the laser arrays can be easily modularized with suitable power sources (nuclear, solar, etc) then they could be placed at various strategic points in the solar systems’ economic sphere to manage fast, return journeys to any suitable destination planet/moon/asteroid. I’ve noted before that Ceres would be an important source of h2O, useful as a propellant. or electrolyzed to H2 fuel/propellant and O2 for life support. With a low escape velocity and slow enough rotation allowing the laser to maintain the beam for the needed time on the craft, this might facilitate this asteroid as a very suitable source of volatiles for commerce and colonization. Given the performance of this craft, this might be a very useful technology where rapid transit is important.
For the shorter-term future, a laser array on the Moon is even more ideal. It permanently faces the Earth with minor libration. There are no weather issues to disrupt the beam. Reducing the acceleration to a comfortable 1g allows the ship to accelerate for 2 hrs, decelerate for 2 hours and reach the lunar orbit in just 4 hours, all under constant artificial gravity. (Whether this can be done depends on whether the beam can maintain that acceleration out to nearly 200,000 km) The craft could orbit the Moon or dock with a space station there, and a shuttle ferry the passengers to the surface. A 4-hour trip is like a current air flight, which is very doable for passengers and logistics. With acceleration and deceleration providing artificial gravity, the flight could be quite comfortable, except for the micro-g start and end of the flight. And no special zero-g toilet with lengthy instructions to amuse the passengers. If not, then a higher acceleration (max 3g) at the start and end of the journey (1 hour each) with a micro-g coast (of 1 hour) for a total of 3 hours, will be less comfortable but tolerable. (Read those zero-g toilet instructions carefully!) It is the sort of performance that the Aries-1B Moon ship has in 2001: ASO, although the LTR cannot land on the Moon and the passengers will need to disembark to a ferry.
But it would make the 3 days Apollo and SpaceX Dragon trips to the Moon seem very primitive and slow by comparison.
Alex,
Thanks for your feedback on our design.
As noted in the paper, the renders that you see are for the “outward-bound burn only, single-use” option. We discussed this issue in Section 3.3 of our paper:
“Active control mechanisms for the reflector would play a critical role in a fully operational laser-thermal propulsion system. In addition to stabilizing the reflector, orientation mechanisms and secondary optics would be necessary to allow the spacecraft to decouple its thrust direction from the incoming beam direction. The design presented in Figures 2 and 9 clearly presents the working principles of laser-thermal propulsion but is limited to thrusting along the incoming laser’s direction.”
In your previous post, you (i.e., Alex) already noted some possible solutions: Repositioning the reflector, gimbling the nozzle, or even having two different exhaust ports.
Repositioning the reflector should be relatively straightforward. It has a mass of just 40 kg and experiences only minor loads. The radiation pressure of the laser flux being reflected is negligible. In our design, the only loading we found necessary to consider was the reflector’s own self-weight while undergoing the gentle acceleration during the laser-powered burn.
While I haven’t done a detailed analysis, I would favor a track or crawler that repositions the reflector along the spacecraft. Robots that can craw along a spacecraft for precision placement of components is something my aerospace colleagues in Canada know something about!
Finally, I would note that the mass of the reflector (40 kg) is so completely negligible in the overall mass of the design that, even if we needed to have two reflectors (one looking aft for the outward burn and one looking forward for the burn-back maneuver), it would not affect the design or mass budgets at all. We have enough margin in the mass budget to add several reflectors, if needed.
Beamed energy propulsion could help push Space Solar Power and vice versa. Senator Shelby-who used to be a Democrat-could replace Manchin’s vote on a Space Solar based build back better…with SLS that has both Union and industry support…as Gene Meyers explained on his most recent Space Show visit. His political contacts will be as vital to you as any engineering. I think you have that down-but Musk is a summer fling. Spaceflight-if it is to survive-needs to get into the ENERGY SECTOR and space solar power can do that. For now, I ask that you table Mars and push space solar that could get widespread support.
“Build back Worse”? Space solar satellite power is very expensive compared to other energy systems, especially ground-based solar and wind. It does have niche applications where cost is not a key factor, such as for military field bases during operations.
But to suggest that this should be done by using the SLS is the height of economic madness, and logistics too. It is very expensive to launch materiel to space and the possible launch cadence for the mass needed would be so low as to make the ISS construction look speedy.
It is spending pork for no value to most of the US population and serves only to keep a small population employed. This is the very problem Eisenhower warned about in 1958.
Large facilities in GEO make very tempting targets in times of war. We may even see cyber disruption of our power facilities in the current emerging conflict with Russia and Germany and Europe are going to feel the cold if that happens.
If we are to build SPSs at all, for beaming power to extraterrestrial facilities and craft, it should be done as economically as possible. That rules out the SLS as a launcher.
I am talking Senators here. I see the writing on the wall. Here is how it goes. SLS is the hook..as Elon hates space solar. We ween off SLS…Starship is bought once working. The point is to do an end run around Manchin. Don’t be a purist.
Why would Shelby vote for the BBB package by offering him more spending on SLS? He represents one of the poorest states in the US and yet votes against the very things that would help his constituents. IMO, it might be much cheaper to carve out $1bn to directly bribe/payoff Manchin and Sinema to vote in full alignment with the Democratic party proposals for the rest of their terms and then retire. Yes, it is banana republic politics, but it is just formalizing the US political funding system and putting the paying for legislation out in the open.
If solar power satellites are to be built, I would prefer it to be done by competitive bidding for a pilot version and then have it become a subsidized commercial operation, just as wind and solar were supported to where they are now cost-competitive with fossil fuels and still getting cheaper, with far fewer environmental costs. Cost of materials to orbit (GEO) from Earth or the Moon is going to be critical and that means the lowest cost launchers must be part of that equation. Maybe O’Neill’s suggested lunar material sourcing and magnetic launcher is the cheapest solution for scaling up the program. IDK. But let’s make sure that whatever the mission solution, it is the right economic choice, not some wasteful boondoggle that serves no one other than those few gaining at taxpayer expense.
Yep, it’s uncanny how … feasible building a laser thermal propulsion infrastructure seems to be. Canada to take the lead in the space race? I wouldn’t mind that at all, eh.
At a minimum, the Martian base will always have a good supply of the highest quality maple syrup. ;)
But no borscht.
What a horrible month.
Perhaps what should be considered is FIRST getting a laser power array inserted into high Mars orbit, with slow chemical rockets. Just to make the Mars aero-braking of future arrivals less risky. Enough to slow the craft, I dunno…. by 25%, so a 100 mega-watt array is not needed…. and does not need to punch through an atmosphere….. maybe a 10 mega-watt (or less?) array beaming for a longer duration than 2 hours. Solar powered with batteries….. sounds kinda too heavy…. maybe multiple launches. Or….. a single, small fission reactor.
How’s it going to be powered?
If you start mass producing solar power satellites, it would be relatively easy for one to reach Mars, just being equipped with ion thrusters to use the energy it produced along the way.
That’s an even better suggestion than mine!
Not sure we will ever LAUNCH from Earth a small, fission reactor (well… the enriched fuel actually) no matter how low the real inherent risk is. The public just wont allow it.
Maybe, if we look around…. we will find fissionable material on Luna (not likely…. unless perhaps an asteroid impact crater….) or Mars.
Neither Orion, nor Daedalus for sure, BUT, it’s a LOT closer to reality! Great design, very inspiring!!
Thanks for the positive feedback!
I wondered why I found this one so familiar, and then realised it’s a variety of the Solar moth concept, which in turn had a predecessor in a design by Krafft Arnold Ehricke. The laser allows for a much smaller reflector, but the principle is quite the same.
https://vimeo.com/256398888
Yes, thanks for giving me the opportunity to elaborate on this point. Solar-thermal propulsion and the Solar Moth in particular were clear inspirations for our design; you’ll find an extensive review of work on inflatable solar reflectors in our paper.
However, there are some important distinctions to make: Solar-thermal propulsion—at least in near-Earth space—uses the solar flux at 1.4 kW/m2. Collecting and concentrating this flux with large reflectors or Fresnel lenses is possible, but the specific power density remains low. The parameter of interest is the “alpha”, expressed in kg/kW of the power system; the lower the alpha, the better. For solar-thermal propulsion designs, it has proven difficult to get the alpha below a value of 1 kg/kW, the value needed to perform rapid transit missions not feasible with chemical propulsion. Also note that it is impossible to concentrate sunlight to temperatures greater than the effective blackbody temperature of the sun; this is a fundamental limit imposed by the Second Law of Thermodynamics and means that solar thermal propulsion cannot have a specific impulse greater than about 2000 seconds.
In comparing solar-thermal and solar-electric propulsion (like Dawn), solar electric is usually going to win out: In both cases, you have a weak but continuous power source. Under this constraint, the highest specific impulse is going to win out, and thus solar electric is usually the winner between these two.
Our laser-thermal propulsion design is quite a different issue: Our laser flux arriving is measured in the megawatt per square meter, not kilowatts, so the effective alpha value is so low (on the order of 0.002 kg/kW), it doesn’t even matter anymore. There is no thermodynamic limit to how much the laser can be focused to achieve high temperatures, and specific impulses of 3000 seconds or greater are possible.
For an intense, short duration burn as found in our design, you actually do not want the highest specific impulse, but rather a combination of high thrust and decent specific impulse. This is what laser-thermal propulsion can deliver. The difference between laser-thermal and solar-thermal propulsion is as stark as the difference between a laser lightsail and a solar sail.
There is one application where solar thermal might offer significant advantages: a solar Oberth maneuver (or “sundiver” mission). When very close to the sun, the greater flux of sunlight can be put to good use. In Section 4 of our paper (Alternative Missions), we speculate that there could be an interesting synergy with our laser-thermal design and a solar-thermal Oberth maneuver: You could use laser thermal at launch from Earth to cancel out much of Earth’s orbital velocity, then fall inward to very near the sun to use the same hardware (inflatable reflector and hydrogen heating chamber) to power a solar-thermal Oberth burn.
Being in a “power rich” design space opens a lot of interesting design possibilities, and hopefully leads to further recognition of the type of unique missions directed energy can enable.
Thank you for your comprehensive reply Andrew. Much appreciated. =)
One example of Ehricke’s solar thermal designs for Convair in 1959.
Convair Solar Powered Spaceship: 1959
He had done several designs, using the inflated “bubble” collector as the core of the concept. At a time when solar PV arrays were not invented, he was very keen on using solar energy to power spacecraft, even using arrays of small concentrators to generate the power for electric engines such as his “butterfly” concept. Von Braun’s wheel space station designs had an annular collector on the wheel to concentrate solar energy for power. This design was used in the movie “Conquest of Space”. [ 1-D curved mirrors to heat oil were often used for solar thermal power before the power tower approach superseded them. ]
While lightweight foil mirrors are attractive from a mass perspective, they do suffer from curvature imperfections, wrinkles, etc. Their shape is maintained by various means, e.g. annular inflated forms around the edge. One can get a sense of this by looking at inflated toy helium balloons that were common a few years ago until the helium shortage reduced their use. The effect of beaming MWs of power to them has not been established. The paper suggests that extra deformations are not a serious problem, but I think that needs to be confirmed by experiment, rather than just computation. The sail designs for Breakthrough Starshot might provide some answers too. For flat sheets of reflective film, origami is one solution to create an even curvature, although the folds may be an issue both for deployment and their effect on the power beam. One wonders if manufacturing thin-film curved mirrors in space is the best answer in the future. It might be a useful technology for both power and imaging. [As collectors it would extend the useful range of solar PV. O’Neill proposed using vast conical collectors to ensure terrestrial illumination levels in his Island 3 colonies up to 3 light days away.]
Hi Andrew
Just how different is a down directed re-entry trajectory? About 9 times Martian gravity directed down to keep a constant altitude initially.
Hi Adam,
The aerocapture design in our study did use down-directed lift to maintain a near constant altitude above the Martian surface, right at the sweet spot where we can—simultaneously—tolerate the heat load (barely) and tolerate the g-load (just barely). As I wrote in the post, it’s gonna be a wild ride.
Getting to share the aerobraking sequence from “2010: The Year We Make Contact” with my Gen-Z students was one of the highlights of working on this project!
Cheers,
Andrew
Nice to see a relatively near-term proposal with a direct path to an interstellar travel design. My guess is that this technology would have a much bigger impact on human space activity than any realistic interstellar travel proposal.
Of course, military applications would drive development of such powerful lasers. That is a problem as the technology would be tightly controlled and lead to arms races amongst the leading technological powers.
Back to the design – the center of gravity needs to aligned with the thrust vector to achieve straight-line acceleration but the drawing suggests otherwise unless the nozzle is slightly canted. Also, can the vehicle itself sustain exposure to the laser light as may occur from time to time? Could such a powerful beam transit the atmosphere without losing collimation? This is a big problem apparently with current directed-energy weapons.
“The 14 km/s Delta-V laser thermal burn sends the spacecraft to Mars on a nearly straight line trajectory”
To get to Mars fast requiring shortening the distance to Mars, rather traveling about 300 million km distance in a Hohmann, this appears to be going about 150 million km and rocket power is used to change the vector of 30 km/sec Earth orbital velocity around the sun.
Or if added 14 km/sec to 30 km/sec, you get 44 km/sec, you get solar escape trajectory [if not intersecting any other body, you will leave our system [eventually].
If added 11 km/sec to hohman transfer 30 + 11], you eventually return to Earth distance. But if doing “nearly straight line to Mars” you aren’t adding the 14 to 30, you added less than 14 {a lot less because not the most efficient trajectory to Mars, it’s shortest distance but require more delta-v.
So I would say this “14 km/s Delta-V laser thermal burn” if didn’t encounter Mars is not doing solar escape trajectory and it doesn’t return to Earth distance like a hohman transfer, would.
Instead it fall back towards the sun and return closer to sun than Earth’s orbital distance- somewhere around Mercury distance from the Sun.
This is excellent work. It updates this concept taking into account advances in optics and lasers. It also applies the technology to current mission needs instead of the original emphasis on getting to LEO. Are there some publications available on this work?
Hi Jim,
Thanks for your feedback. Our design study has now been published (Duplay et al. “Design of a rapid transit to Mars mission using laser-thermal propulsion,” Acta Astro. 192:143-156 (2022)). You can get a web browser friendly version of the paper here: https://ar5iv.org/html/2201.00244
If you were asking about the advances in optics and lasers that would enable a very large phased-array, there is much active work ongoing in this area. In addition to Philip Lubin’s group at UC Santa Barbara, the Breakthrough Starshot initiative is supporting other groups working on scaling-up phased arrays, including at Australian National University:
Chathura P. Bandutunga, Paul G. Sibley, Michael J. Ireland, and Robert L. Ward, “Photonic solution to phase sensing and control for light-based interstellar propulsion,” J. Opt. Soc. Am. B 38:1477-1486 (2021). http://dx.doi.org/10.1364/JOSAB.414593
While I’m not an expert on photonics (i.e., we simply took this technology as a given in our design study), my understanding is that no fundamental showstoppers to building arbitrarily large phased arrays have been identified. Cost likely remains the main obstacle to building a 100-MW-class laser array.
As I alluded to in Part 1 of my post, I like to think of Arthur Kantrowitz as my “academic great grandfather”: I did my MS and PhD in Abe Hertzberg’s Aerospace and Energetics Research Program at UW-Seattle in the 1990s, and Hertzberg was, in turn, Kantrowitz’s former student. I recall you (Jim Early) knew Hertzberg as well.
It’s gratifying that the photonics revolution of the last two decades is enabling us to come back to Kantrowitz, Kare, and Hertzberg’s concepts for directed energy propulsion; just unfortunate they are not still around to see their ideas bearing fruit.
I like the idea that the fuel pipes are around the outside of the combustion chamber which would result in the same cooling of the combustion chamber. The only problem I have with this laser thermal rocket idea is that I don’t think this technology would result in a 45 trip to Mars considering the low specific impulse which would be more like four months instead only one and a half months. It would still need a lot of fuel like a nuclear thermal rocket.
Don’t nuclear thermal rockets get to Mars in about 60 days and don’t have a large delta-v difference at Mars?
It seems if laser thermal rocket as same or more mass of nuclear thermal rocket, one should probably just use the nuclear thermal rocket.
This paper gives a ISP of 3,000 which is the same as a gas core nuclear thermal rocket. It says on Google online that it could take as little as four months. VASIMR can go to Mars in only 39 days, but only with a high ISP, but one article says it needs 200 MW of nuclear battery reactors as Alex Tolley has written.
Seems almost too good to be true (which is why I am highly suspicious). On the surface it doesn’t seem like there’s any problems here which properly arranging laser arrays couldn’t solve in some detail as to perform transfers to wherever you wish to go. The only problem I do see here is the fact that given the limitations of power transfer and the time required to achieve a particular final velocity you may have to be carrying a lot of initial mass as propellant. The ability to expend this propellant meaningfully might be not as simple as portrayed here. But I guess it least it has the advantage of having your energy source somewhere else.
Giving you the benefit of the doubt I wonder whether or not this laser arrangement could be coupled with the Bussard interstellar Drive in a new fashion than it has been proposed up to this point. Consider this: one moves a constant velocity with your collection field extended to gather up interstellar hydrogen which you place in storage aboard the ship; this is being done at the same time you’re expending propellant provide the heating of the onboard hydrogen that you already have with a interstellar laser shining on the craft. If your intake is greater than your outflow you end up filling the tanks and can then engage in a acceleration maneuver again under laser power which will increase the speed of your ship up to the point where you require more additional gathering of interstellar hydrogen, to begin the refueling process of your vessel. Has anyone considered this as a possible way to circumvent the nuclear reaction limitations that the current Bussard interstellar Drive has ??
I’m wondering if the reflector could be made of solar sail material that has a far lower aerial density. Silvered for the reflecting side, unsilvered from the canopy. If all of the mirror could be made this way, it might mass 2% of the reference mirror. This would allow it to have a larger area, reducing the heat load per unit area, and also allowing for the phased light lobe to be useful to a longer range, a longer, but lower acceleration time, and lower laser power. The tradeoff might be the required longer focal length and hence added mass of the truss supporting the engine, not to mention more difficulty controlling the mirror to maintain its focus on the engine. Regardless, as we know that Breakthrough Starshot is expecting to use GW lasers on tiny foil sails, it should be possible to reduce the mass of the mirror and/or increase its size to improve performance.
The reference design mission architecture is interesting as the payload is released to make the journey, and the main craft quickly returns to earth orbit to be refueled, rather like a staged launcher. If refueling can be done fairly quickly, perhaps robotically with automated docking at a fuel depot, then payloads could be fired at Mars (or other bodies) in rapid succession, perhaps hours or days apart, allowing for an effective multiple payload mass to be delivered to the destination. It might be possible for the craft to deliver successive payloads throughout the year, simply adjusting the delta_V to the needed orbital trajectory. This might be more suitable for a Mars settlement than infrequent, but much larger payloads on Hohmann orbits every 2 years using conventional rocketry.
“Robert Zubrin often makes the point that—even if advanced propulsion capable of high thrust and high specific impulse was available—he would still opt for a 6-month free-return trajectory and use the enhanced propulsion capability to bring more payload.”
From that perspective, remember that cargo is also shielding, if arranged right. Wouldn’t astronauts headed to Mars prefer to arrive with an abundance of supplies and equipment?
These ‘hot’ trajectories just strike me as wasteful, for someplace as close as Mars. They make more sense for destinations in the outer Solar system, like Jupiter or Saturn, where a minimum energy trajectory would consume a significant fraction of a person’s life, and the launch windows are relatively infrequent.
Zubrin dislikes any contrary opinion to his. His approach is to land a working return rocket first that makes the fuel for the return trip ensuring the astronauts following on have a guaranteed trip home. Similarly, having all the base supplies sent on ahead makes the same sense. What you want for the crew is to have as rapid a trip as possible to reduce fatigue, adaptation to micro-g, psychological stress, systems malfunctions, etc, etc. Studies on simulated Mars flights have demonstrated the psychological problems that develop over the trip time. It is no different to people on Earth preferring to fly to a destination than to take a slow trip by sea or rail to reach their destination.
As the LTR can make more trips than a Hohmann trajectory, there is more opportunity to schedule the outward-bound flight. Where Zubrin has a point is that a failure to make a Mars rendezvous, a fast trajectory is harder to recover from than a free return trajectory back to Earth.
The point about easier return in case of some failure is strong. The first missions to Mars with humans onboard have to be slow. Definitely even in our times some plucky guys could be found who would be OK with wild rides while still qualifying for the mission, but if they fail and perish in the attempt, this would discourage others for long…
Agreed. But note this LTR proposal is for 1 MT payloads, not crewed vehicles. A rapid supply of needed supplies makes sense, especially if a critical component that cannot be supplied locally is needed. With a fuel depot in orbit, multiple payloads could be directed at Mars in short order, allowing for redundancy in case of failures.
I know I would be somewhat concerned about being shot towards Mars in a capsule that required exquisite aerobraking maneuvers to make a safe Marsfall. [c.f. the scene of Jupiter aerobraking in the movie 2010: Odyssey 2]. There needs to be some backup method to mitigate potential miscalculations with crewed vehicles. Capture facilities in Mars orbit would be nice in some future situations. At least in the short term, however, it supports my belief that robot exploration is the way to go. If the cost of such transportation is low, transport is fast, and robots expendable (until laws on sentient AIs end that), then the logic of using such machines becomes even stronger. Any human settlement of the solar system will have robots paving the way first. For the outer system, we will need very fast ships if humans are to migrate there, or travel will be in O’Neill’s with engines to make the journey comfortable and no more onerous than spaceship Earth’s journey through space.
“From that perspective, remember that cargo is also shielding, if arranged right. Wouldn’t astronauts headed to Mars prefer to arrive with an abundance of supplies and equipment?”
Have return and abundance of supplies and equipment already at Mars orbit. Using efficient Hohmann Earth to Mars, but for crew, I think a requirement should be 3 months or less with the immediate crew abort option to return to Earth.
Though could be returning to Earth via Venus.
It seems in low radiation environment- of Solar Max- and if already have established crewed base, one could do the 6 months to Mars, but for Mars exploration program, one should have 3 months or less, option.
While analogous focused solar thermal propulsion concepts are not new, and chamber temperatures are hard-limited by the temperature of the Sun, they definitely have some pros over lasers. The energy source is available everywhere – it is the Sun, and it means very much.
-No need for risky aerocapture at target
-No need for high thrust at the acceleration leg. This means also much less requirements for chamber design since pressures and fluxes are many times lower
-Much more available power for other spacecraft systems
-With inflatable solar concentrator, some solar power could be available in the outer Solar System, maybe even in the Kuiper Belt.
-Development of this design also brings solar-thermal-powered solar Oberth maneuvres much closer.
While Isp is lower, using lithium hydride at 3500 K still gives 700-800 s, maybe even higher if not only exhaust is fully dissociated but lithium is also ionized.
On the other hand, high-power lasers are available only on Earth for the foreseeable future, and focusing is limited by atmospheric turbulence. 1 arcsecond of divergence already means 250 m beam width at 50000 km, and it is likely to get worse with increased flux because air will heat in the beam and it is much more difficult for adaptive optics to correct (because of positive feedbacks which are absent in the systems used for observation)
PS I realized that closed-cycle solar thermal power generators also benefit from the same type of solar concentrator. If the working fluid is heated in the described chamber but directed into a turbine, a very high efficiency is allowed by Carnot equation because of great available temperature difference between “hot” and “cold” sides. This can be used for arcjets and other power-demanding applications in space far from Earth.
Solar powered thermal rockets (STR) have much lower Isp, require much larger mirrors, and more propellant.
I ran some numbers before my most recent comment above, and it is quite clear that a small mirror with beamed energy has a far better overall performance than a STR. The mirror alone, even if made of sail material would still mass more, and that is apart from the problems associated with such a material. With lower Isp, the propellant becomes a far greater fraction of the craft, so that the Mo/Mi ratio of the rocket equation becomes far larger for the same delta_V.
While I suspect this reference design needs some tweaking and further development, I cannot fault the concept and general numbers as a reasonable optimization of the performance given the stated mission goal defined by NASA. What will be an issue is the 100MW phased laser array. I suspect that this is likely to be something Lubin pushes as part of his DE-STAR roadmap (and Breakthrough Starshot) and this craft piggybacks on that infrastructure when in place (by the DoD?).
Conceptually, I really like this idea.* I hope it proves to have legs and that a demonstration version (CubeSat scale?) proves the concept valid, and the full scale (and larger) craft are built as the needed phased laser arrays are built.
* I am biased in favor of such approaches
While lasers could provide more focused flux, it quickly runs against capabilities of chamber materials. And this is aggravated by the need of high chamber pressure with laser launch, because of short acceleration legs. On the other hand, the conventional thermal load limits, 3000-4000 K, are achievable with the focused solar thermal heating, too. If hydrogen can be stored cryogenically, then the same chamber temperature amounts to 1500 s because H2 atomises at these temperatures and low pressures. That’s 10 km/s for payload-to-propellant ratios close to 1:1. For propellant heating, reflector does not need to be in precise parabolic shape, it just has to achieve spot size comparable to the size of the image of the sun _if_ the reflector had precise shape. The hard-limit for flux at the focus is on the order of flux at the surface of the sun, and it is still 60 MW/m2. The closer to the Sun, the less are requirements for shape. I’m more concerned about the mass offset, the solar reflector has to be stabilized against torque and other forces, as it was said in the comments above. But this, too, is mitigated by much gentler accelerations since the leg is not limited by laser range.
My guess is that in the absense of nuclear power in space and lasers that could outshine the Sun everywhere in the Solar System (Breakthrough Starshot ultimate ones would do), solar concentrators coupled with some kind of high power electric generators and high-trust electric propulsion are the ultimate concept of this kind. The above-mentioned 200 MW VASIMR could be powered by 1 GW of sunlight, which could be collected by a kilometer-sized reflector. If it is made from 1 um aluminized polyimide, it would weigh several tons, which does not seem way too much for big martian missions. Inflating it to 0.001 Pa would add some 10 kg of gas while still providing the same several tons of total inflation force over the whole surface. Shape maintenance calculations are entirely different level, of course, but there is one strong point here – areal density of space-based solar collectors could be brought 2 or even 3 orders of magnitude less than that of all current or near-future solar cells. For high solar power in space, the benefits of any kind of solar concentrators are great.
The lasers, although, are the way to Breakthrough Starshot, and they enable many other ways to explore Solar System. Maybe even beam-ride launch from ground to LEO. Water heated to 3500 K still gives Isp quite suitable for ground-to-LEO, and imagine how the cost per kg would drop if there is no need for cryogenical storage, red-ox mixing, and the propellant itself costs nothing! Considering energy and power, if we have multi-gigawatt CW-lasers, then surely there are ways to launch things from ground to space without having to burn fossil fuels, at least onboard :-)
PS In no way I’m saying I don’t like lasers. What I argue most, is that atmospheric turbulence is a no-go for ground-based BS lasers and they have to be build in space. But the same beams in space, multi-gigawatts in single milliarcseconds, are completely another deal. They outshine the Sun everywhere. They could message anyone in the Galaxy, who is not obscured by dust. They could sublimate ices from Kuiper belt bodies, enabling experimental studies in outer Solar System. They could illuminate Planet Nine, to see it better after it is found. They could produce ablation thrust on incoming Earth impactors from millions of kilometers (spot size 50 m from 1e6 km for 10 mas beam, 1.6 megawatt per square meter for 4 GW, that heats to 2300 K and easily boils rocks)…