Solar Thermal Rocket

[Moonguy]

Dude, every time I read one of your posts I get at least 100 ideas to chew on. I'm starting to mutter things like "Inverse square rule at Saturn would yield. . ." at the dinner table. My family is starting to worry. STOP!! Get out of my head!!!

[Enthalpy]

The Solar thermal rocket enables an extensive mission around Saturn. This first part describes up to the capture by a first moon.
http://en.wikipedia.org/wiki/Saturn
http://en.wikipedia.org/wiki/Moons_of_Saturn
http://solarsystem.nasa.gov/planets/profile.cfm?Object=Saturn

A Delta IV Heavy shall place 8808kg at 3162m/s (10km²/s²) above Earth's gravity. The eighteen 4.57m Solar engines add 7176m/s, leaving 4944kg heading to Saturn in 5+ years. Acceleration can take more than 12 days, so lighter engines are better, with diaphragms or a small secondary mirror to limit the power at one Sun-Earth distance. At Saturn's 9.582UA, each engine pushes 29mN and uses 0.2kg/day hydrogen.

The necessary inclination of the capture orbit at Saturn already constraints the launch window, so no gravitational assistance is taken from Jupiter, but smarter people could use Earth and possibly Venus to save on the launcher or extend the probe; both are welcome for this high-energy mission.

Outer moons have too inclined orbits, so the first target is the puzzling two-tone Iapetus.
http://en.wikipedia.org/wiki/Iapetus_(moon)
http://solarsystem.nasa.gov/planets/profile.cfm?Object=Sat_Iapetus&Display=Facts&System=Metric
Preferring Nasa over Wiki, the orbit has 3561Mm radius and 8.313° inclination versus Saturn's equator; lower moons are equatorial.

The probe arrives with 5760m/s above Saturn's gravity and brakes by 4596m/s (502 days) before a 4195m/s pass at 3813Mm from Saturn's center. 276m/s more braking (30 days) reach a capture orbit of 14490Mm*3561Mm covered in 320 days - make science meanwhile. Two peri-Saturn kicks totalling 872m/s (95 days) achieve the circular 3561Mm orbit covered in 79.3 days.
Well, at least if believing my spreadsheet: SaturnWeakBrake.zip

Gravitational assistance by Iapetus (escape 573m/s) or an other moon might help.

The probe reaches a polar Iapetus orbit to study it, with periapsis 1.2Mm from Iapetus' center (the moon has 735km radius) and apoapsis 15Mm in Iapetus' orbital plane (Iapetus' influence reaches 36Mm against Saturn). Computed from 200m/s above Iapetus' gravity, the manoeuvre passes with 342m/s at 1.92Mm from the moon's center, and reaches the final orbit in a single 169+23m/s (11.9+1.6 days) brake consuming 192m/s.

IapetusWeakBrake.zip

Smarter people might enter the moon's gravity where Saturn's one makes it easier - if any feasible.

The described capture by Saturn then the moon consumes 5744-8=5736m/s, leaving 3116kg orbiting Iapetus.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Hohmann from Iapetus to Hyperion (orbit R=1501Mm at 5027m/s in 21.3 days) would take years, so the transfer shall push continuously. The probe brakes by 1763m/s and loses the orbit inclination of 8.31° or 601m/s as a mean. For this, the engines push up or down by 41.5° for 2*90° in each turn around Saturn, and flat for 2*90°. The perfectible de-tilting costs 253m/s, bringing the spiral transfer to 2016m/s.

Beginning the push in elliptic orbit around Iapetus gives free 8m/s.

Hyperion's influence reaches 2.2Mm against Saturn. The target orbit is polar 0.5Mm * 1.1Mm covered in 2.6 days, with periapsis in Hyperion's orbital plane at the daylight side. Capture computed from 50m/s above Hyperion's gravity passes with 34m/s at 0.60Mm from the moon's center, and reaches the final orbit in a single 2.0+0.4 days brake consuming 34+6m/s: 10m/s are for free.
HyperionWeakBrake.zip

From 3116kg in Iapetus orbit, the 2016-8-10m/s transfer puts 2653kg in Hyperion orbit in 129 days.

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Around the following moons, orbits computed without Saturn's influence have a period similar to the moon's orbit around the planet, and their apoapsis exceed the moon's reach against Saturn's gravity gradient, so my capture costs are inaccurate.

To obtain an estimate nevertheless, which should be pessimistic, I separate the transfer between the moons' orbits from the capture by the moon, computed with the weak thrust, from zero speed above the moon's gravity, and neglecting Saturn. This would nearly be the case if the major axis were polar; though undesireable for the exploration, it protects the apoapsis from Saturn's tides.

It is my hope that specialists find better scenarios. Maybe the probe can enter the Moon's well near the outer Lagrange point being temporarily the apoapsis; the first orbit around the moon would then have its apoapsis perpendicular to Saturn hence protected, giving time to lower it enough. A polar orbit with periapsis in sunlight over the moon's equator should be possible.

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From Hyperion to Titan (orbit 1222Mm at 5571m/s in 16 days), the spiral transfer takes 544m/s in 31 days, as Hohmann saves nothing for near orbits. But gravity assists at Titan?

While Hyperion offers 10m/s, capture at Titan (3.1Mm * 25Mm) by weak thrust is counted as 298m/s. From 2653kg in Hyperion orbit, the 544-10+298m/s transfer puts 2481kg in Titan orbit in some 100 days.

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2421kg leave Titan after 298m/s expenses. Big jump from Titan to Rhea (orbit 527Mm at 8483m/s in 4.5 days). Six engines are discarded to save 150kg; the 12 remaining push 0.35N and eject 3.4kg/day.

The spiral transfer begins with 2271kg, costs 2912m/s and 198 days, ends with 1796kg. Hohmann would save 120m/s but is too long. Slingshots at Titan, maybe Rhea?

Capture (1.6Mm * 11.9Mm in 3.3 days) is to cost 107m/s and puts 1780kg in Rhea orbit.

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1764kg leave Rhea for Dione's orbit (377Mm at 10030m/s in 2.7 days), but of course the probe visits Dione's trojans: Helene before and Polydeuces after - same orbit, just separated by 60° from Dione.

The 1547m/s spiral transfer takes 86 days and leaves 1557kg.

Helene's (R~18km) Lagrange points are 71km away, not bad for observation if the instruments see that near. Or a minimally elliptic and tilted Saturn orbit might make a single-turn helix around Helene's orbit. Negligible fuel expense.

Covering the 60° to Dione in 60 days costs 2*25m/s (would be more at Titan's orbit). Capture at Dione (1.1Mm * 7.3Mm in 2.3 days) costs 86m/s and puts 1540kg in Dione orbit.

Leaving Dione and joining Polydeuces puts 1523kg there. The R~1.3km object can be sniffed from all directions at negligible cost. Its mass and density are still unknown; to my taste, estimated 200µm/s² gravity and 0.7m/s escape speed are slightly too much to land there and analyse probes, but unknown smaller moons may float near Helene and Polydeuces.

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1523kg leave Polydeuces for Tethys' orbit (295Mm at 11339m/s in 1.9 days), and the probe visits also Telesto and Calypso, Tethys' trojans.

The 1309m/s spiral transfer takes 63 days and leaves 1371kg. The Lagrange points are ~38km from the trojans, no fuel expended there.

60° in 60 days cost 2*20m/s. Capture at Tethys (1.1Mm * 4.8Mm in 1.8 day) costs 77m/s and puts 1357kg in Tethys orbit.

The probe weighs 1343kg at Calypso.

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1343kg leave Calypso for Encelade's orbit (238Mm at 12623m/s in 1.4 days). No trojans are reported, so it's time to check the places.

The 1284m/s spiral transfer takes 55 days and leaves 1211kg. Observe new moons there - maybe.

60° in 60 days cost 2*16m/s or 3kg to join Encelade. Encelade's Lagrange point is only 950km from its center, and the surface 252km... So the probe should not orbit Encelade, but rather follow a minimally elliptic and tilted Saturn orbit that makes a single-turn helix around Encelade's orbit.

I budget 50m/s or 5kg to stabilize around Encelade, putting 1203kg there.

I would not separate a lander nor diver there. This is worth an own mission. But a penetrator maybe. And spend all due time there, for sure.

The probe weighs 1195kg at the other Trojan place. Find new moons maybe, possibly land on them to analyse probes. A spring lets take off.

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More to come.
Marc Schaefer, aka Enthalpy

[Enthalpy]

Leaving Encelade's orbit, the probe keeps six engines and discards six to weigh 1045kg and join Pallene (212Mm at 13375m/s in 1.15 days).

The 752m/s spiral transfer takes 52 days and leaves 983kg. Search for unknown moons on the path - the spiral transfer eases it. The probe has probably optical and radar imagers to inspect the known moons; a lidar may also help discover new ones. Occasionally, a concentrator for an engine, for electricity production or for radio transmissions can pump a properly doped Yag with 200W Sunlight.

The R=2.5km Pallene costs nothing to approach and leave, but is too big to land.

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983kg leave Pallene for Methone (194Mm at 13982m/s in 1.01 days).

The 607m/s spiral transfer takes 39 days and leaves 936kg.

Methone has R~1.6km so I'd rather land elsewhere. But near passes permit to fire bullets (hydrogen cannon?) to help analyze the surface, and maybe evaporate some material with the pulsed Yag.

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936kg leave Methone for Mimas' orbit (185Mm at 14301m/s in 0.942 days). No trojans are reported, so it's time to check the places.

The 319m/s spiral transfer takes 20 days and leaves 912kg at the first Trojan place.

60° in 60 days cost 2*12.5m/s or 2kg.

Mimas' (R=198km) influence extends only 520km against Saturn, so the probe can't properly orbit the moon; I consider again a slightly tilted elliptic Saturn orbit that circles Mimas. For that, I budget 50m/s or 4kg at arrival, and as much at departure.

One more hop to visit the other Trojan place, and the probe weighs 900kg.

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The 746m/s spiral transfer to Aegaeon (168Mm at 15047m/s in 0.808 days) needs 44 days and leave 847kg. Detect small moons and rocks meanwhile, since a ring begins there, and land to analyze them.

With R~250m, ~20µm/s² gravity (1/10th the probe's thrust) and 0.1m/s escape speed (spring), approaching Aegeon and landing costs no propellant.

200W concentrated Sunlight can evaporate snow at some distance for analysis, including without landing, but the hotter Yag is better; Sunlight may be good for an oven. Hydrogen at engine exhaust has also enough enthalpy per mole, as kinetic energy, to evaporate many materials at a moon from a limited distance; replacing temporarily hydrogen with, say, rubidium or xenon, enhances sputtering at the target, and special chamber and nozzle can keep the jet more concentrated.

Three engines are kept and three discarded to weigh 772kg. The big hydrogen tank should already have been discarded, keeping a small one, but I've forgotten it.

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The 1411m/s spiral transfer to the main A ring (140Mm at 16458m/s) needs 138 days and leaves 689kg. As usual, detect small moons and rocks meanwhile, and land to analyze them.

Venus and Earth flybys, plus smarter transfers and captures at Saturn's moons, would leave much more probe mass. The Solar thermal engine enables the described dumb scenario, but smart methods developed for chemical and ion engines apply here.

At the ring, capture the smaller rocks and land on the bigger to analyze them. Continue to dive as long as hydrogen suffices.

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Saturn's main rings can be a separate mission. The probe passes nearer to Saturn to save capture propellant, brakes at Titan and changes the orbit inclination there for flexible launch opportunities; orbiting no moon save further propellant. Ejecting vapour instead of hydrogen, the probe finds propellant in the rings and can navigate everywhere without limit.

Marc Schaefer, aka Enthalpy

[Enthalpy]

As Nasa calls for bids for a mission to Europa, Jupiter's moon supposed to have an ocean of liquid water, I should like to remember the script I proposed in post #25

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry769456

based on the Solar thermal engine whose current description begins with the present thread

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/

 

The existing script is absolutely straightforward and quick, directly from Earth to capture by Jupiter then Europa - something enabled by the isp=1267s from the Solar thermal engine. Though, gravity assist would bring a probe bigger than 900kg to Europa: slingshot at Venus, possibly Earth and Ganymede. I won't explore this scenario soon.

[Moonguy]

This occurred to me while I was reading the earlier posts: Would it not be possible to channel the concentrated light into the thrust chamber through fiber-optic cables? This would make the orientation of the reflectors independent of the thrust vector. The light from the fiber-optics concentrates inside the chamber as in other concepts, but does not require the chamber to be designed with exotic materials in order to allow light through. I have not had time to check it out, but it would seem this could scale sufficiently for use by a manned vehicle. . . especially one going to Mercury (or Venus).

[Enthalpy]

Thanks for your interest!

 

The thrust can be fully oriented independently of Sun's direction by using secondary mirrors, which will probably be present for optical reasons anyway. I suggested a setup there

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/#entry752817

but it's merely to suggest that vectoring is possible and rather easy. An optics designer would improve that.

 

I have of course nothing against fibres! Just that last time I threw thoughts at them, what is possible and what isn't were unobvious. I suppose they won't replace the concentrator. As well, the vast collecting area is a design constraint and I wouldn't like to waste even 30% of it, a usual filling factor for fibre bunches.

- It's easier at Mercury, sure...

- The very efficient use of Solar power converted to kinetic energy is a huge advantage of the Solar thermal engine. Ionic engines for instance would outperform the ejection speed and thus save gas mass, but 30% light to electricity and little converted to kinetic energy makes their Solar panels really impractical for a significant thrust. A nuclear reactor making electricity would require a bigger radiator than the Sunlight concentrators here, while a reactor heating hydrogen wouldn't reach the 2800K for hydrogen dissociation that enable the 1267s isp.

 

In the manned Mars mission script, I need thrust like 100N, and the concentrator area is still feasible (...not small) for the Solar thermal engine.

 

My design for the heater doesn't need transparent materials. It's just tungsten that absorbs light at the vacuum side and heats hydrogen at the other (sure, light at limited incidence and gas at fins)

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/#entry753432

"Focus" there designates an immaterial location, the entrance of a hole full of vacuum.

Tungsten would be exotic in a car or a PC. For a launcher or satellite, tungsten and its machining are less exotic than niobium alloys for instance.

 

ESA had tried to design such an engine, but had a window and an exchanger made of a rotating bed of ceramic pellets. These were serious feasibility worries and performance limits (isp=800s) at their design. By removing these difficulties, my design looks feasible, and even rather easy. I believe it can become a big thing in space transport, not just a Mercury, Mars or Europa, but also for the geosynchronous orbit.

 

The ruminator and some regenerator details are nice features. As usual, I didn't check if some Sapiens had invented them before.

 

What I'm less pleased with are the concentrators. Made of electrolytic nickel (plus coatings) like satellite dish antennas, they should weigh like 1kg/m²: less would be useful. Though, I wouldn't like to increase them beyond the fairing's diameter. D=4.57m is easy to store and deploy even in big numbers, and the engines can be tested on Earth.

[Moonguy]

Good clarifications, thanks! Regarding the reflector/concentrator, why are they made of anything rigid at all? Solar sail films already exist with 10 GRAM/m2 density. They ;also have 90% reflectivity. The materials available - at least the samples I have held in my hands - are strong enough to be integrated into an inflatable structure or even a mechanical arrangement like that in an umbrella. Adding scrim or other items to strengthen the reflector material is easily accommodated well with in the 1 kg/m2 mass limit.

 

The system you describe has a lot of flexibility. I have to think a solar thermal stage would be easier to develop than a nuclear thermal stage. The higher Isp (1267 vs ~950) sure looks like ample incentive. Sort of gets you to wondering why. . .

Edited May 20, 2014 by Moonguy

[Enthalpy]

If someone can propose a film as the concentrator, I have absolutely nothing against, sure!

 

Sunlight must be concentrated from a D=4.57m reflector to a focus like d=20mm. While heating does not need an optical quality reflector, it does need a resulting Sun disk nearly the minimum size - if not, the too big hydrogen heater would lose more power through its radiation.

 

I couldn't think of a film setup accurate enough, hence the ultra-thin nickel with stiffeners, as for dish antennas. Deployable film reflectors exist for radio dish antennas, but their wavelength is less demanding. I also like that 4.57m concentrators (or 10m launched by the SLS) allow to test individual engines on Earth in existing vacuum chambers.

 

The heater has about the same 2800K as the filament of a light bulb and radiates almost as effectively... It loses only a fraction of the incoming light power because the Sun is even hotter - but for this comparison to hold, the thermal design must be optimized. That's the aim of my reflective-coated regenerator insulated by vacuum and of my ruminator.

 

As well, optimum operation at Jupiter makes an imperfect heater at Earth, for instance. Either have two specialized heaters, or optimize for the more demanding location, and cut-out the excess sunlight. For Jovian missions, I found the acceleration near Earth quick enough.

 

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A nuclear thermal engine heats and expels hydrogen just like my Solar thermal engine does. The Solar has born advantages to achieve a higher Isp:

 

- It can use just tungsten because no neutronic properties are required. Hotter hydrogen is faster.

- At 2800K and only 30mbar, 23% of the injected H₂ split into H, absorbing much more heat from a limited temperature. During the expansion, a good fraction of this heat converts to kinetic energy. This was not considered in ESA's design. A reactor would also try to push strongly, and the corresponding chamber pressure prevents hydrogen dissociation.

 

The pressure is constrained by the expansion possibility. In my present design, the mean free path at the exhaust is a few time smaller than the nozzle diameter, so hydrogen behaves like a gas. Less expansion would waste heat not converted to kinetic energy, while more chamber pressure would reduce the dissociation and the isp. I had failed this estimation previously (2010) when I got isp~2200s.

 

A "boosted" operation mode, where more hydrogen gets heated by as much sunlight, increases much the thrust and wastes little isp. It seems useful during certain operations, but at least for the trips around all equatorial Jupiter or Saturn moons, I found no advantage to it. For a capture by Mars or Mercury maybe, since one m/s is worth more near to the planet than far away.

 

I too believe that my engine is rather easy to develop, but as an old engineer, I know well enough that a convincing system design is just the requisite before running into the worries of development.

[Enthalpy]

The sunlight concentrators can fulfill more purposes at Europa or elsewhere:

• As the reflector of a high-gain datacomm antenna.
• To concentrate sunlight on an electricity generator:
  http://saposjoint.net/Forum/viewtopic.php?f=66&t=2051
• To pump with sunlight a laser, maybe Cr:Nd:YAG. The laser beam might evaporate points of Europa's surface for remote analysis.
• To pump a laser that transmits data. Consider my filter at the receiver at Earth
  http://www.scienceforums.net/topic/74445-evanescent-wave-optical-filter/

Marc Schaefer, aka Enthalpy

====================================================================================

Here's how a chemical engine and the Solar thermal one can share the performance to escape a planetary orbit and achieve an asymptotic speed, with more details than
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/#entry754921

A chemical engine can provide a brief kick near a planet which brings much asymptotic speed according to Oberth
http://en.wikipedia.org/wiki/Oberth_effect
the optimum share is when a relative variation of the begin-to-end mass provides the same speed variation from both engines, including the multiplying factor by Oberth effect.



i_sp=465s for the hydrogen+oxygen RL-10B and heirs, i_sp=392s for a kerosene+oxygen engine for GTO and escape missions
http://www.scienceforums.net/topic/73571-rocket-engine-with-electric-pumps/page-2#entry772828
give an optimum share at 4298m/s (18.3km²/s²) and 3530m/s above Earth's gravity, 1934m/s and 1589m/s above Mars', but:

• This neglects all inert masses. Two-stage launchers prefer less.
• This optimum is very broad. 1934+500m/s from Mars loses 0.4% mass overall, 1934+1000m/s 1.7% - room for other constraints or preferences.

Marc Schaefer, aka Enthalpy

Edited May 24, 2014 by Enthalpy

[Enthalpy]

When pumping lasers with sunlight, say at a Europa mission, my filter might also direct to the lasing ledium only the useful wavelengths, to limit heating; it must be inserted where light is parallel.

A good 200mm mirror would concentrate light to a 100mm spot from 20km altitude - if only the pulsed laser produced a perfect beam. 150W mean optical power evaporate Europa's surface thanks to the short pulses.

If analyzing the shallow surface isn't enough, one could build a hydrogen gun on the space probe, to shoot dense bullets like 1g at 3km/s. These will penetrate deeper in ice, and the probe can shoot 10,000 of them of varied composition to map the moon.

Marc Schaefer, aka Enthalpy

[Enthalpy]

The continuous faint push of the solar thermal engine makes bad use of the Oberth effect
http://en.wikipedia.org/wiki/Oberth_effect
Pushing near a planet is much more efficient, so that a strong short kick (possibly at slower exhaust speed) can be advantageous.

The solar thermal engine can be operated with a higher propellant throughput for that, accepting less hydrogen dissociation and a lower temperature. It brings a modest advantage.

An other option, combinable with the previous one, is to release some heat not obtained in real time from sunlight.

For instance molybdenum absorbs 375kJ/kg to melt at 2896K, niobium 288kJ/kg at 2750K, or hafnium... Ceramics or salts may bring more options. How tungsten resists these is very unclear to me. As much heat as 1000s from a D=4.57m concentrator takes 60kg molybdenum near Earth, but just 2.2kg near Jupiter, where the usefulness is clearer.

Or add heat from electric current, as in a resistojet. A safe Li-polymer battery store 475kJ/kg, other chemistries more - and better, spacecraft have already a battery. Electricity from sunlight is less efficient, so the continuous push better results from sunlight-to-heat, but lightweight and shared storage is attractive for kicks, especially near Jupiter or Saturn.

Marc Schaefer, aka Enthalpy

[Enthalpy]

I suggested there how to share the job among a chemical engine that lets a craft escape a planet and solar ones that accelerates it further

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry807522
This improves further if the efficient solar engines first raise the apoapsis and the strong chemical engine uses the Oberth effect to escape only - or when arriving at a planet, capture chemically to an eccentric orbit, then circularize by the solar engines.

This is an optimum without inert masses, for this more general elliptic orbit:

 

At Earth for instance: an Atlas V 505, an H-IIB or an Ariane 5 can put 18.8t on a 400km orbit (Ariane 5 Me can more... Wider fairing please!). 15 D=4.572m solar engines bring 14503kg to a 2-day orbit with 127300km apogee; this takes 9 months, good for a space probe. An RL-10 kicks 11450kg to asymptotic 4233m/s - or a lighter and more efficient 15kN engine possibly with electric pumps, more careful with the payload, usable at the destination as well.

For 2902m/s performance to the elliptic orbit, the solar engines (at 90% isp because of long pushes) bring 1.46* more mass than an RL-10; the advantage increases because the inert mass of a two-stage launcher penalizes an escape mission; and the combined scheme also saves mass at the destination.

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At telluric bodies, probes shall usually land or go to low orbit, and sunlight suffices for the engines. In some cases, the chemical kick gives the whole transfer speed.

 

We can send these masses by the improved scheme - now we're getting somewhere.

 

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At Jupiter and Saturn, the final orbit depends on the mission, which must give time to the solar engines.

I decribed there a mission to all of Jupiter's equatorial moons

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/#entry756421
instead of multiplying all masses by 0.78 following the wrong estimation of the launcher's capability, 7932kg in transfer permit to multiply them by 2.05, so the mission ends at Io with 800kg instead of 304kg.

The mission to all of Saturn's equatorial moons, launched by an Atlas V 505 instead of Delta Heavy,

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry773439
ends at the A ring with 980kg instead of 689kg.

The transfers need such speeds if no flyby helps - optimum to Jupiter and Saturn, accelerated to Uranus and Neptune:

 

The improved scheme sends these masses - in transit for Jupiter and Saturn because the missions vary too much from there on:

If a mission needs short pushes there, a hydrogen resistojet and a battery may perform better than direct heating by sunlight.

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At Uranus and Neptune, the solar engines would take too long. Chemical engines impose then eccentric orbits.

The masses in the table are for twin missions in direct transfers, more accelerated to Uranus, similar to

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry757109

http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry757109

(multiply all masses there by 0.78: then 600kg in orbit, the new scheme improves to 2000kg)

The probe arrives at the point that puts its apside line in the planet's equatorial plane and brakes near the planet. This may put the periapsis in daylight or not. In a first phase, the probe observes the planet from there and looks for new moons. Then the inclination and the apoapsis are reduced in small steps to overfly the irregular non-equatorial moons as possible. When in equatorial orbit, the probe observes the lower moons.

Marc Schaefer, aka Enthalpy

[Enthalpy]

A mission to Jupiter's moon Europa, where wild scenarios put a water ocean below the ice crust
http://en.wikipedia.org/wiki/Europa_(moon)
benefits from the improved role sharing among the solar and chemical engines, but weaker sunlight there needs time.

The already described improved Earth departure sends 7932kg to Jupiter - though I'd have nothing against Venus, Venus and Earth flybys.

The probe arrives with asymptotic 5643m/s. The chemical engine brakes by 803+618m/s at 670.9Mm from Jupiter's center - Europa's orbit radius - where the escape speed is 19431m/s. This leaves 5810kg with 10000Mm apoapsis nearly in Europa's orbital plane.

The solar engines lower the apoapsis from 5073m/s to 541m/s over Europa's (approximately) circular orbit. Eighteen D=4.572m concentrators push for 3 days around the periapsis so the process takes around 3 years; make science meanwhile. The long push is 80% efficient, so 3683kg head to Europa.

Gravitational assistance by Jupiter's four big moons should have brought observations and saved time and mass, but isn't easily accessible to hand computation.

The chemical engine brakes by 73+144m/s at 100km over Europa's surface for capture, or 1661km from the center, where the escape speed is 1963m/s. This puts the apoapsis at 10Mm, below the Lagrangian distance, with the preferred orbit inclination, leaving 3512kg.

The solar engines circularize the orbit. They provide 431m/s in three months, leaving 3392kg at 100km over Europa, better than 899kg previously
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry769456
This includes all the propulsion (about 800kg). The mass can comprise an orbiter, a lander and a diver.

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The concentrators have uses beyond propulsion, at Europa and elsewhere, not only as a radiocomm antenna or to warm the spacecraft.

Each concentrator catches 827W, or together 15kW. Filters can direct the best frequency band to each use; consider my evanescent wave filter
http://www.scienceforums.net/topic/74445-evanescent-wave-optical-filter/



Concentrated filtered light can pump a laser for data transmission and an other to analyze the surface of a celestial body (besides a hydrogen gun). Solar cells get efficient with strong light that doesn't heat them unnecessarily; varied semiconductors would get each the best frequency band. 40% conversion from small cells would already provide 6kW electricity, nice for a radar and for radiocomms. A turbine would deliver more
http://saposjoint.net/Forum/viewtopic.php?f=66&t=2051

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More moons deserve their own space probe, like Saturn's Enceladus.
http://en.wikipedia.org/wiki/Enceladus_(moon)
Weaker sunlight there would make the circularization really long, but some answers exist.

• The concentrators can make electricity over the whole orbit, stored in a big battery, used at periapsis in the solar engines in a hydrogen resistojet mode. Not very good at Jupiter, but interesting at Saturn.
• The solar engines can trade efficiency for strength.
• More concentrators, or even (gasp) a lighter probe.
• Moon flybys are probably the key to success. The solar engines bring the fine orbit correction and the capacity to adjust the periapsis and the inclination.

Marc Schaefer, aka Enthalpy

 

[Enthalpy]

It goes without saying, but maybe better if I say it: concentrated sunlight, split over wavelength bands, has uses not only at Europa. Here electricity production for a geosynchronous satellite:



As usual, the satellite's body revolve once a day around the North-South axis to face the Earth, and at its North and South faces, some parts revolve once a year to face the Sun - but here the solar cells can turn with the body.

• It needs no slipring to transmit the electricity, only a Nasmyth path or similar.
• Solar cells of varied bandgap can be separated. No difficult new technology, and we can have more different materials. That's much cheaper and efficient. They get their best band through filters.
• The smaller cell area saves costly semiconductor and work. Concentrated light improves the efficiency further. Heat stays bearable thanks to the good conversion. Fluid cooling is an option.
• The unconverted light, typically mid-infrared, may power a turbine, or be dumped.
• The concentrators should be lighter than solar panels of same area, and more so of same power.

At 50% efficiency, two D=4.57m concentrators (to fit flat in the fairing) provide 22kW electricity.

Marc Schaefer, aka Enthalpy

[magma time light]

Electro

Magnetic

Regeneration

or as I like to think, flying saucers.

[Enthalpy]

The sunheat engine can bring a decommissioned Keyhole-11 from Earth orbit to Lunar orbit to make better images there.

As an evolving family of spy satellites, the KH11 is little known
https://en.wikipedia.org/wiki/KH-11_Kennan
and much data is inferred from the Hubble Space Telescope, supposed to use the same D=2.4m primary mirror and chassis
https://en.wikipedia.org/wiki/Hubble_Space_Telescope
so 240nrd resolution would separate 75mm at 300km distance from the Moon too, better than presently 500mm with the Lunar Reconnaissance Orbiter
https://en.wikipedia.org/wiki/Lunar_Reconnaissance_Orbiter
and improved datacomms would provide us clearer images from more sites.

Secrecy is a hurdle, especially since the added module provides data storage and transmission, needing to know at least KH11's downlink format - but maybe it can be reprogrammed. The optical and mechanical data is a smaller worry. On the other hand, carrying the toy away avoids to deorbit it, and the saved propellant gives an operational life extension worth 100M$: incentive. And did I read that the Nro had already offered decommissioned orbital KH?

Many orbservation and spy satellites orbit the Earth: Landsat, Spot, Helios, Lacrosse... I take KH11 as an extreme example, as most others are lighter. Some are also less secret. Or Nasa can use the 2.4m mirrors given by Nro to build a new, light Lunar craft carried there from Leo by the sunheat engine; though, I'd prefer to send these to Mars, an operation I may describe later.

---------- Mass, speed and propellant

An Atlas V 541 puts a 13t propulsion and data module on the same Earth orbit as the chosen KH11, say 300km x 900km x 98°. The module deploys its sunheat engines, navigates to the KH11, and grasps it delicately for 40N push. I hope no electric contact is needed: the KH11 keeps its energy supply and transmissions, the module intercepts, processes, stores and retransmits the data.

I estimate the empty KH-11 weighs 13t, for no good reason, and the aggregate starts with 26t.

 

1. Perigee raised to 400km costs 28m/s. This action overlaps with the beginning of the next one.
2. Apogee raised to 326Mm, the Moon's Lagrange point's distance, costs 2931m/s. The aggregate weighs now 19955kg.
3. Perigee raised to 326Mm and inclination changed from 98° to almost Lunar orbital plane (1022m/s forward, 89m/s polar). Cost 1026m/s leaves 18373kg.
4. The Moon is there at that moment and catches the aggregate in a 1737+300km x 58Mm polar orbit. Aposelene sinked to 300km costs 291m/s and leaves 17901kg on the 300km x 300km Lunar polar orbit: that's 4901kg more than the KH11, for the module's structure, engines, propellant rest and data tinkering.

The 15 sunheat engines push 2.7N each. This permits Hohmann transfers, making them as good as the ion engine that offers more Isp but must spiral due to its fainter thrust. As the kicks extend for long around the periapsis, I take mean 90% efficiency for the 12424m/s ejection speed at steps 1, 2 and 4. The whole transfer takes approximately 15 months.

---------- Structure

The D=4.572m concentrators travel stacked horizontally. Each weighs 25kg.

The 115m³ tank for 8099kg hydrogen is mean 4.4m wide and 8.5m tall. Welded AA7020 tubes machined to L=1.18m Ri=32mm Ro=33.1mm make its hexagonal truss (200kg) with 12 nodes per turn while AA7022 sheets machined to e=1.3mm except at the seams make the skin (500kg), welded at the truss for the cylindrical part and unsupported at the heads. The truss could be twice finer-grained and the skin have integral ribs.

30mm foam (180kg) let 37kW leak in. Over 15 min launch operations, the hydrogen warms by 0.6K.

50 plies of 13µm Mli (250kg) let 14W leak in, evacuated by a ~600Wm cryocooler.

At each tank end, a truss of aramid or glass fibre holds to the tank. With 12 nodes per turn and 2 stages each, using L=1.18m Ri=35mm Ro=37mm tubes, they weigh 180kg together and leak <2.5W together.

1310kg insulated and supported tank and 375kg engines leave 3200kg for the equipment bay and optional hydrogen to manoeuvre in orbit.

---------- Data processing, storage and transmission

Let's take 30,000 pixels wide images for KH11, with 2*12bits colours. At 75mm resolution and 1554m/s it must read 20700 lines/s, slower than at Earth: either the module averages several lines, or the KH doesn't tilt its view when observing - too little is known here. This produces 1.9GB/s, faster than can be transmitted.

5000 chips of 32GB Slc Flash store 24h worth of uncompressed sensor data. Easy to increase.

On Lunar orbit, 6 concentrators are reassigned to solar cells of varied bandgaps through filters, as suggested here on Sep 20, 2014. 40% conversion provide 53kWe on dayside, of which 20kW=+73dBm transmit data. The battery for 53kW*68min weighs 470kg.

6 other concentrators work as transmission antennas. Each gains 62dBi at 30GHz - choose a frequency with muuuch available bandwidth - or 70dBi phased together. 3 own ground stations for 24h coverage have 15m collectors gaining 73dBi. After 3dB propagation losses, they receive -21dBm - the receiver can consist of a preamp, an oscillator and two diodes towards the complex baseband.

If 100MHz are available and the noise temperature is 30K=-184dBm/Hz, the signal-to-noise is +83dB. 6dB margin permit 150M symbols/s of 23 bits each, for decoded 300MB/s. Phase noise isn't trivial, and the encoding shall suppress low frequencies. Can the reprogrammed KH-11 transmit so quickly?

We shall modulate the polarization too, for a 3D-constellation. This was demonstrated over fibres, maybe it's already done with radio. Throughput *1.5 or 450MB/s. Still less than the sensors. 1000MHz available bandwidth would have multiplied the throughput by 8.7. Simultaneous beacons to many smaller separated ground stations would have improved too.

Transmitting 450MB/s half of the time needs 4 years to map all the Moon with 75mm resolution uncompressed. Storage on Earth takes a few thousand disks.

Marc Schaefer, aka Enthalpy

[pavelcherepan]

Great idea. Really like reading your posts, but can't usually comment on anything because you put a lot of thought behind what you write.

 

Still, I wonder whether there is a significant enough necessity to map the Moon at 75 mm resolution? Wouldn't it be better to concentrate resources on preparations for manned Mars mission?

[Enthalpy]

Thanks for your interest!

 

So do I wonder... About the Moon, I believe to understand that only a small fraction has been imaged with 0.5m resolution by LRO because transmissions limit the amount of data, so a complete and detailed mapping would be an improvement, be it at 0.5m or 75mm. Also, as Esa proposes to Nasa to settle a base on the Moon's remote side, having good pictures first would be advantageous.

 

Mars, sure! This was my first intent with KH-11. What stopped me: apparently Altas V Heavy doesn't start from Vandenberg, so I couldn't get enough mass on a sun-synchronous (polar) orbit, and I redirected my effort to the Moon. This explains some features oversized for the Moon, especially the datacomms. Though, the description for a lunar mission is a good start for the same at Mars.

 

Concentrate the resources on manned Mars: yes... I consider my sunheat engine is essential for it

http://www.scienceforums.net/topic/83289-manned-mars-mission/

and the sunheat engine must be proven first on an automatic mission. Transporting an existing KH11 to the Moon needs little more than developing the engine, hence is a cheap test. This delicately megalomaniac undertaking is also a convincing illustration of the engine's possibilities. 18t is about what a Saturn-V launch put on Lunar orbit.

[Enthalpy]

Service Pack 1 for the Keyhole-11 to the Moon...

----------

The orbit's inclination must be changed before raising the perigee: 1120m/s. Sinking the aposelene costs 605m/s. This leaves 17274kg on Lunar orbit, or 627kg less. More hydrogen also increases the tank and insulations by 111kg, so the ~3200kg equipment bay loses 738kg. Still easy.

The module but fits in the Delta long 5m fairing now. Though, the equipment bay is still heavier than needed, so the module doesn't need to exploit the full capacity of the Delta V 541 (even a 531 should suffice) so the tank shrinks. If needed, the Delta operator claims it can expand the fairing.

----------

Transmissions to several Earth stations simultaneously, less oversized and hopefully less wrong now...

Two 2m*5m solar panels suffice and are easier to orient. 30% efficient, eclipsed 50% of the time, they provide mean 4kW, of which 2kW feed the transmitters. That's pessimistic because when eclipsed by the Moon, the craft doesn't transmit.

Here 13 small ground stations at sites with usually clear weather are separated by >2400km to cumulate the throughput of several beams - better than exaggerating a constellation modulation:

6 stations are generally visible from the Moon, sometimes 5, briefly 4. Allowing for maintenance and cyclones, the craft sends 5 beams to chosen stations. Hard disk drives shipped by boat can replace high-speed cables.

Each of the craft's D=4.572m 50GHz antennas has its first zero excentered by 600km at Earth's equator, so optimized sidelobes are weak at 2400km, and the beams well separated. Aperture synthesis could replace five reflectors but is uneasy at 6mm wavelength; distinct amplifiers for the beams, possibly to distinct and interleaved antenna elements, would then preserve the amplifiers' efficiency. I take five reflectors here.

Each beam modulates two polarizations and phases in a 4D (not 3) constellation, 4D-cubic for simplicity, hence equivalent to 4 independent amplitude-modulated channels. Each of the 5*4 channels gets mean 100W supply, so 40% efficiency at full power lets it radiate 80W=+49dBm for the full-power symbols.

Free-space loss is 227dB at 384Mm distance and 6mm wavelength, emission by 80%*D4.572m gains 66dBi, reception by 80%*D4m gains 65dBi, 3dB are lost, so the receiver gets -51dBm for full-power symbols of each polarization and phase.

100MHz available band permit 40MBaud; smooth symbol transition put the noise bandwidth at 40MHz. 30K noise temperature make -108dBm for each polarization and phase. With 7dB margin, the full-power symbols are 50dB stronger than the noise, or 316* the voltage.

256 points per polarization and phase between -316 et +316 noise voltage are separated by 2.48 noise voltages, so a symbol has 453ppm chances per polarization and phase of being wrongly received (no Viterbi nor treillis here, provided it exists in 4D). An RS(255,243) code to correct 6 symbols has then 5e-11 chances of being wrong, or one 100Mb image in 0.4 million.

5 beams at 40MBaud of 4 bytes per symbol and a 243/255 efficient code transmit, if the Moon eclipses the craft half of the time, 0.38GB/s - almost the sensor's throughput as its swathes overlap. From our R=1737km Moon, 26PB for a D=75mm resolution full map is transmitted in 2.2 years.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Reassigned concentrators help to transmit data from Martian orbit to Earth. The probe can only target one ground station.

At 1.524AU (Sun-Earth distances), seven D=4.572 dishes concentrate 64kW light on solar cells of varied bandgaps. Collective 50% efficiency and 50% eclipse time by Mars leave mean 16kW supply to feed the transmitter half of the time with 7000We for each polarisation and phase. Again a 4D constellation, cubic for simplicity, with amplifiers 40% efficient at full power, transmits 5600W=+67.5dBm per polarisation and phase at the strongest symbols.

Seven D=4.572*80% in-phase dishes radiate at 50GHz with +74dBi gain. At mean 1.823AU=273Gm, free-space losses are 284dB. A D=24m*80% reception dish gains +80dBi. After 3dB losses, the ground station receives -66dBm per polarisation and phase at the strongest symbols.

100MHz available band permit 40MBaud and the noise bandwidth shall be 40MHz. 30K noise temperature make -108dBm for each polarization and phase. With 6dB margin, the full-power symbols are +36.4dB stronger than the noise, so each polarization and phase swings from -66 to +66 times the noise voltage.

64 points per polarization and phase are separated by 2.10 noise voltages, so a symbol has 0.30% chances per polarization and phase of being wrongly received. A RS(4095,3935) code on 12b symbols, which have 0.60% chances of being wrong, corrects 80 positions. A few sockets of 20-core Xeon decode it using Log and Exp tables, or better one Xeon Phi.

4*6b per symbol at 40MBaud and the 3935/4095 efficient code transmit, if Mars eclipses the beam half of the time, mean 58MB/s. Of R=3390km Mars, 14PB for a D=0.2m resolution full bicolour map is transmitted in 7.5 years.

Minimum data compression improves that. Or if 200MHz are available, 80MBaud of 4*5b transmit mean 96MB/s. Faster, or more colours, or better resolution.

Marc Schaefer, aka Enthalpy

[Enthalpy]

This is how to bring a decommisioned WorldView from Earth orbit to Mars orbit. We couldn't transmit a complete map at full KH-11 resolution anyway. As opposed, the lighter WorldView 2 and 3 add eight colour bands to the panchromatic one, including true colours that are all-important for public support.
http://www.satimagingcorp.com/satellite-sensors/worldview-2/
http://www.satpalda.com/product/worldview-2/
https://en.wikipedia.org/wiki/WorldView-2
WorldView-2 nears its end of life, WorldView-3 leaves some years more, and GeoEye is a similar candidate, all with image quality as you experience on Google Earth. The Spot series has a worse resolution, Helios maybe.

WorldView-2 moves with 7469m/s at 773km altitude. I hope to put it at sun-synchronous 350km above Mars so 3384m/s there give the same line frequency. This improves the panchromatic resolution from 0.46m to 0.21m and the multispectral one from 1.85m to 0.84m but shrinks the swathe from 16.4km to 7.4km. I checked that a D=1m F=3m paraboloid is perfect within 1nm at 773km and 2.5nm at 350km once the focal plane is adjusted - but what happens to the probable Ritchey-Chrétien is parsecs beyond my skills.

----------

The transport combines chemical propulsion at escape and capture to exploit the Oberth effect with my sunheat engines as I described here on Jul 27, 2014. This table is only from Earth to Mars hence should be clearer:

 

An Atlas V 511 or Delta IV M+(5,2) launches the 8100kg tug to 773km 98° - or an Ariane V, an H-II, maybe a Falcon 9 if building light. The tug grasps the 2300kg empty WorldView. From 10400kg, the transport leaves 4610kg on low Martian orbit, or 2310kg for the used tug.

The engine is a throttled down RL-10, possibly without nozzle extension. Since 8kN thrust would suffice, it could be a 1bar pressure-fed or 70bar electrically pumped design.

10 sunheat engines take about 10 months at Earth and at Mars.

A slingshot at our Moon could spare the chemical engine. Capture at Mars by the sunheat engine remains less efficient, but the craft is simpler.

----------

To the KH-11 design, the tug adds the RL-10, a small oxygen tank (toroid on the sketch) and is smaller.

 

 

Scaled like the hydrogen volume, the tank and truss would weigh 670kg, ten sunheat engines 250kg, the RL-10 300kg, leaving 1t for the equipment bay.

----------

At Mars, two engines control the attitude and orbit, four supply electricity through solar cells of varied bandgaps, four make the already described transmission. The 4/7 weaker received field loses one bit, from 6 to 5 per polarisation and phase, transmitting mean 45MB/s including the eclipses. The 0.21m 11bit complete map transmits in 4.1 years uncompressed, and the 8 colour channels with 4*4 times less resolution in 2.0 years.

Marc Schaefer, aka Enthalpy

[Enthalpy]

I suggested sister crafts orbiting Uranus and Neptune:
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/#entry756556
whose mass were inaccurate but improved anyway by my better scenario
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry818683
leaving 2t at destination.

Even at Neptune, 30.1AU=4.50Tm from the Sun, concentrators can direct the 1.51W/m² sunlight on solar cells. Fifteen D=4.572m concentrators and 50% efficient multigap cells would produce skinny 176W electricity, but the unfoldable AstroMesh claims to be bigger
http://www.northropgrumman.com/businessventures/astroaerospace/products/pages/astromesh.aspx
up to D=25m for AM-1 and D=50m for AM-2, of which D=12m has flown. D=25m would provide 352W electricity and D=50m 1408W without plutonium.

Their RF reflector is a mesh but I hope some metallized film would reflect sunlight. The manufacturer claims fuzzy ~0.3kg/m² improving with size, or 150kg for D=25m. The fuzzy shape accuracy and maximum frequency include 26GHz.

I wish the Astromesh would concentrate sunlight for my sunheat engine, but I suppose it's not accurate enough. Targeting solar cells is easier.

Earth is close to the Sun as seen from Uranus and Neptune, so the same concentrator would double as an antenna, for instance with one secondary mirror made of mesh to redirect only the RF and some mechanical or electronics means to steer the RF or light independently.

Modulating the phase and the polarization cumulates again the bandwidth, but with the wideband noise exceeding the signal from Neptune, Hadamard and Reed-Muller soft-decoding gains only as the Log of the band spreading.

Marc Schaefer, aka Enthalpy

[Enthalpy]

EDRS is a new relay satellite in geosynchronous orbit that receives data by laser from Earth-observation satellites (typically on low orbit hence seeing the ground stations shortly) and relays it by radio to ground stations
http://www.bbc.com/news/science-environment-35446894

I suggested here on Jan 10 and Jan 15, 2016 to transmit several beacons to separate ground stations and encode data on the polarization as well
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-3#entry900362
and this works from 36,000km distance 100 times better than from the Moon to transmit much more than EDRS' 180MB/s. Several primary sources targeting different ground stations can share one mirror.

Lasers pumped directly by sunlight as I suggested here on May 24, 2014 may be more efficient
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/page-2#entry806581

Marc Schaefer, aka Enthalpy

Edited January 31, 2016 by Enthalpy

[Enthalpy]

Esa tries to involve Nasa in a manned base there, and my sunheat engine can transport heavy freight to the Moon. The comparison with other engines holds for other missions.

An Atlas V or Ariane 5 shall put 16.0t in Earth orbit, just 400km high, approximately in the Moon's orbital plane. 28.5° inclination is good.

---------- Leo to Lmo with sunheat engines

• Successive kicks at the 400km perigee raise the apogee to 326Mm. This means 2931m/s, but the long kicks are about 90% efficient.
• A few kicks raise the perigee to 326Mm and adjust the inclination. 1026m/s, efficient.
• The Moon passes by. The craft brakes at 300km periselene to be captured and to lower the aposelene from 58Mm to 300km. This means 605m/s at 90% efficiency, plus 50m/s adjustments.

The transfer needs 4955m/s performance. I_SP=12424m/s let the craft weigh 10.7t on a 300km*300km Lunar orbit.

Ten D=4.57m engines make the operation in 15 months, 10 of them for perigee kicks. I neglected the eclipses; launching at the best season void nearly avoid them at perigee.

Hydrogen is used fast and often enough to avoid a cryocooler. Taking gas or liquid adjusts the tank pressure.
http://www.scienceforums.net/topic/60359-extruded-rocket-structure/page-2#entry761740

---------- Lmo to lunar surface

Deorbiting from 300km costs 63m/s, braking before the surface costs 1880m/s, hovering 3*20s over potential sites and hopping 2*200m cost 160m/s
http://www.scienceforums.net/topic/85103-mission-to-bring-back-moon-samples/#entry826166
for a total of 2103m/s.

The main and attitude engines can burn at 100bar Pmdpta+O₂ expanded in D=1.2m to 1.4kPa for 50kN and I_SP=3763m/s=384s. This lands 6.1t on the Moon, including 131kg of Li-polymer batteries that rotate the pumps and are a precious payload too
http://www.scienceforums.net/topic/73571-rocket-engine-with-electric-pumps/#entry734835

Liquid oxygen needs probably a cryocooler, while Pmdpta is storable on the Moon.
http://www.chemicalforums.com/index.php?topic=56069.msg254340#msg254340
To restart from the Moon after a night, all propellant pairs need some thermal control. These don't freeze.

-------- Alternatives to Pmdpta+O₂

MMH and N₂O₄ achieve only I_SP=3519m/s=359s with the same pumps and land 5.9t. They are toxic and can freeze. Pressure-feed would be worse. Time for retirement.

H₂ and O₂ give I_SP=4423m/s=451s and 99kN in a short RL10A, and attitude control thrusters exist already. It would land 6.6t but needs a hydrogen cooler it this technology restarts from the Moon. Fuel cells to power the pumps would land 0.2t more after a significant development
http://www.scienceforums.net/topic/73571-rocket-engine-with-electric-pumps/?do=findComment&comment=737220

-------- Alternatives to the sunheat engines

Starting from 7434kg in 28.5° Gto, one RL10A can reach the Moon, orbit there and land, with insulated tanks. It costs 474m/s to 326Mm apogee, 1026m/s to 326Mm perigee, 605m/s to 300km circular Lmo, 2040m/s to descend and land, totalling 4195m/s. Excellent I_SP=4423m/s=451s land 2.9t, an incentive for the sunheat engine.

Among the electric engines, the Vasimr has good thrust and power efficiency
http://www.adastrarocket.com/aarc/VASIMR
https://en.wikipedia.org/wiki/Variable_Specific_Impulse_Magnetoplasma_Rocket
The VX-200 shall use 200kW electricity to push 5N by ejecting argon at I_SP=49030m/s=5000s. The peak power is huge, so let's have the craft spiral from 400km Leo to circular 326Mm then to 300km Lmo. It costs 7675-1088m/s then 1554-302m/s, totalling 7839m/s. Beginning with 16.0t at Leo, the craft weighs 13.6t at Lmo, seemingly 2.9t better than the sunheat engine. 2364kg argon take 274 days (plus the eclipses) to eject, acceptable.

But let's check its solar panels: at the International Space Station, each 34m*12m Solar Array Wing carries 16400 silicon cells of 80mm*80mm to produce 32.8kW during daylight when new
https://en.wikipedia.org/wiki/Integrated_Truss_Structure#Truss_subsystems
so 200kW need 6 Solar Array Wings. 500µm silicon make then 50% of 0.7t lost from the payload (ATK's UltraFlex claim 150W/m2 or 1333kg), and the wings occupy over 68m*36m and cost 6/8 as much as on the ISS.

Even if accepting some nuclear reactor: 33% conversion into electricity demand to radiate 400kW; at 400K, it takes 2*138m².

The 10 sunheat engines need only 164m² of sunlight concentrators of thin metal.

The sunheat engine needs 7* less watts per newton than the Vasimr, it doesn't waste 4/5 in a conversion to electricity, and its collector area is cheap and light.

Marc Schaefer, aka Enthalpy