Exoscientist

Astronomers want to plant telescopes on the Moon. The lunar surface offers advantages for infrared and radio astronomy, despite the challenges. By Ramin Skibba, Inside Science | Published: Tuesday, January 19, 2021 Astronomers want to plant telescopes on the Moon | Astronomy.com (See the link there to a journal special issue exploring the idea.) I speculated about the possibility of detecting exo-civilizations optically in this Kerbal forum post: How large a space telescope do we need to see exo-civilizations? - Science & Spaceflight - Kerbal Space Program Forums In the discussion in that thread, someone suggested we would need a telescope 1.6 km across to see a visible disk of an Earth-sized planet at the nearest stars. But we might not need to be able to resolve a visible disk to be able to observe illumination of the exoplanet beyond that which would be expected on its nightside. In any case quite large telescopes could be made on the Moon if you used rotating liquid mirrors. This proposal is for one 100 meters across: Texans Want to Put a Big Ol' Liquid Mirror Telescope on the Moon But can they do it? And why even try? BY TIM CHILDERS NOV 18, 2020 https://www.popularmechanics.com/space/moon-mars/a34714863/liquid-mirror-telescope-on-the-moon/ On Earth there are limits to the size you can make a liquid scope because the rotating mirror surface and containment vessel creates wind currents that distorts the liquid mirror surface. But this would not be a problem on the airless Moon. So that raises a question: is there a limit on the size you can make such a mirror on the Moon? Another possibility would be to do the detection through radio telescopes on the Moon. The advantage of radio telescopes is they don't have to have a solid surface but can consist of a set of grid wires, as was done with the Arecibo telescope. And this is the approach taken for one plan for a radio telescope on the Moon: Apr 7, 2020 Lunar Crater Radio Telescope (LCRT) on the Far-Side of the Moon. Saptarshi Bandyopadhyay NASA Jet Propulsion Laboratory https://www.nasa.gov/directorates/spacetech/niac/2020_Phase_I_Phase_II/lunar_crater_radio_telescope/ So how big would a radio telescope have to be on the Moon to detect Earth-like radio emissions from a near-by star like Alpha Centauri? Note this is a different question than that studied for example by SETI. With SETI they assumed such a civilization was beaming radio emissions directed at us. Such searches have been negative. But in the scenario I'm considering, an advanced civilization is creating omnidirectional radio emissions just as a byproduct of conducting its advanced civilization. How large a radio telescope would we need to detect those? Robert Clark

Yes, that is a good question. Here's a NASA page showing the full disk at night of some regions on Earth. The lighting is pretty spotty. https://earth.app.goo.gl/rDX9W5 Bob Clark

Thanks for that. Perhaps we could just observe the variation in light between the day and night side as a point source. Say, for reasonable surface materials for a rocky planet, the night side brightness is too high and would be best explained by artificial illumination. Bob Clark

To be able to directly image an exoplanet the space telescope would not need to be especially large; the James Webb Space Telescope could do it. What would be key though is a starshade to block the stars light. See the discussion here: https://en.wikipedia.org/wiki/New_Worlds_Mission The animation illustrating a proposed starshade is pretty cool on that page. The starshade also would not need to be particularly large, perhaps 10 meters across. But the issue with the nightside city lights is that it likely would be much less than the full daytime reflected light of the planet, unless, the civilization is something like Asimov's Trantor, a single city covering the entire surface of the planet. Bob Clark

A team of scientists is investigating ways of detecting exo-civilizations aside from just radio signals as with SETI: JUNE 19, 2020 Does intelligent life exist on other planets? Technosignatures may hold new clues. by University of Rochester https://phys.org/news/2020-06-intelligent-life-planets-technosignatures-clues.html Two methods of detection mentioned in the article are detection from reflected light from solar panels or detection or pollution such as CFCs. However, even on our planet the number of solar panels would not be such that they would add appreciably to the Earth light. And CFCs presence might be short lived as it has been on our planet, having been banned. Could we instead detect the light on the night side coming from all the artificial lighting that would be used in a civilization? Some of the photos seen from space of the cities alit at night on Earth have been quite striking: Aug. 14, 2014 Space Station Sharper Images of Earth at Night Crowdsourced For Science https://www.nasa.gov/mission_pages/station/research/news/crowdsourcing_night_images How big would a space scope need to be able to see this in the Alpha Centauri system, for example? Robert Clark

Excellent. Thank you very much. Bob Clark

This article suggests you will be able to see the full disk Earthrise from the lunar surface: Earth Rising Earth as seen from the Moon is always in the same place – true or false? It depends. By Paul D. Spudis MAY 15, 2014 https://www.airspacemag.com/daily-planet/earth-rising-180951474/ While most locations on the Moon's surface would only be able to see partial disk Earthrises, at the lunar terminator (separator between far side and near side of Moon) and at the lunar poles, you'll be able to see the full disk Earthrise. Anyone want to do a try in Kerbal Realism mode to show what this would look like? Intriguing question: will we see a "huge Earth illusion" like we see a "huge Moon illusion" when looking at the full disk? MOON ILLUSION IS ALL IN YOUR HEAD BY: BOB KING NOVEMBER 24, 2015 https://skyandtelescope.org/observing/moon-illusion-confusion11252015/ Bob Clark

Right. Let me redo that calculation: Use equation x=1/2 a t2, and round off 1 g acceleration as 10 m/s2. Then 2x1016 m = (1/2)*10*t2. So t = 63x106 s. That's 730 days. The speed reached is v = at =10*63x106 = 630,000 km/s. This is double lightspeed of 300,000 km/s. Then you would have to consider relativistic effects. At some point as you approached light speed you would need to expend a huge amount of energy to maintain that speed, still without actually reaching lightspeed. Perhaps someone will do the calculation about how long it would appear to take for shipboard time compared to Earthtime for the journey. Bob Clark

The time is actually shorter, and it doesn't quite reach light speed. It's surprising how fast and how far you can go by just 1 g acceleration: A nice way to remember the distance of a lightyear is that it's about 10 trillion km, which equals 1x1015 meters. Use equation x=1/2 a t2, and round off 1 g acceleration as 10 m/s2. Then 2x1015 m = (1/2)*10*t2. So t = 2x107. That's 231.5 days. The speed reached is v = at =10*2x107 = 200,000 km/s, less than lightspeed of 300,000 km/s. Bob Clark

I would have loved to have seen a true microscope sent to Mars. After more than half a dozen landers there still has not been sent a true optical microscope sent. The best resolving power has been at no better than that of a geologists hand lens. This would have importance for the search for possible life, but also for geological samples. Bob Clark

Yes. They could use sea level engines for all six engines. You would lose on delta-v or payload though because you wouldn't get the high vacuum Isp of the vacuum engines. By the way, it is not well known that with altitude compensation you improve both sea level thrust and vacuum Isp for the sea level engines. So you would get better takeoff thrust actually than just 6 sea level engines. The vacuum Isp part is well known. But sea level thrust is also improved because sea level engines with fixed nozzles are a compromise. Even for sea level engines a large portion of the flight of that first stage takes place in near vacuum conditions. So to get good performance then also, you use larger than ideal nozzles for sea level. This decreases the sea level thrust. But with altitude compensation you make the nozzles adapt to ideal size both at sea level and vacuum and intermediate altitudes in between. I've been trying to get some Kerbal experts to do the calculation in Realism mode to see what the improvement is in payload or delta-v for engines given altitude compensation. It's not that difficult to do, really. For instance for just a standard fixed nozzle, if you want an accurate simulation you have to include how the thrust, so effective Isp, varies with altitude. Such variation has to be done with altitude compensation also; its just a different formula. The needed formulas are also well known. Just nobody has ever done it. My point is the improvement in payload with altitude compensation is quite high. For instance a rule-of-thumb among propulsion engineers is "every 10% increase in Isp results in 100% increase in payload." The reason of course is because of the exponential nature of the rocket equation. As Ray Kurzweil has noted, people really don't have a good intuitive sense of the nature of exponential growth. Most people would think "OK, so you improve the Isp 10%; so why would I spend this amount of money to increase payload 10%?" The issue is it is much better than 10%. A rough estimate I did was an increase of 25% for a two stage rocket. For a parallel staged rocket like the Falcon Heavy or the SLS, 40%. And for a SSTO, 100%. Because nobody has done an accurate simulation to see how payload is improved with alt.comp., nobody thinks it is worthwhile. Once you open yourself up to the possibility it might be useful it really doesn't take much thought to then come up with various different ways of doing it at low cost. Bob Clark

Elon on Twitter said hopefully the Starship will fly this week:https://www.cnet.com/google-amp/news/el ... -fly-soon/Presumably this will be a short hop test. But running the numbers the Starship could get a surprisingly high delta-v.The latest version has 3 level and 3 vacuum engines. Presumably only the 3 sea level ones will be used at launch. So this will mean 3 x 200 tons = 600 tons of launch thrust. I’ll take the propellant load as 400 tons to allow for payload at later tests. But launching the bare rocket could get 354*9.81Ln(1 + 400/120) = 5,100 m/s delta-v, past Mach 16.This though is the delta-v as an expendable. Some of the propellant has to be used though for landing so the actual delta-v achieved will be less than this.This scenario though illuminates the importance of achieving altitude compensation. Those 3 vacuum engines have to just stay idle during launch, like dead weight. Imagine instead having altitude compensation so all six engines could be used for liftoff. You would have 1,200 tons liftoff thrust. Nearly the full propellant load could be used. You would get in the range of 7,000 m/s delta-v. Actually it would be higher than this since the altitude compensation would also allow you to use the full vacuum Isp of the vacuum raptors of 380 s at high altitude. This would mean quite a large portion of the Earth could be covered by point-to-point rocket travel. This represents a huge market for the Starship.Bob Clark

Odd, that this didn’t occur to me earlier, but this idea for the Starship used as a first stage means it wouldn’t be the stage going to orbit. Then the payload loss for reusability would be less since you wouldn’t have the greater weight for thermal protection for reentry from orbit. In fact, the Falcon 9 booster has minimal thermal protection weight. But this is also due to the fact it slows down to minimize thermal heating on reentry from suborbit. Intriguingly, the payload loss for the Starship used for this purpose may be less than with the F9 because it is higher temperature steel rather than aluminum-alloy like F9, thus requiring lower slowdown and lower fuel burn. But going by the estimate of the F9 for just the booster reusability it would only be 30% loss due from partial reusability. Then the partial reusability payload would be in the range of 50 to 70 tons. Note that New Glenn using partial reusability is about the size of this Starship intermediate-size version and also uses methane fuel and gets about 45 tons to LEO. So likely this Starship version would likely beat the New Glenn in payload and speed to active flights. Note as well the Starship might be able to then get close to the original 25 to 1 mass ratio without the extra thermal protection an orbital vehicle would require. Now comes the controversial part: Spoiler Alert! If you don’t like considering ideas not accepted by the majority, don’t read below. Spoiler If the Starship really is able to get the 25 to 1 mass ratio without the reusability systems, then running the rocket equation it could get ca. 45 tons to LEO as an expendable SSTO. This is significant payload, so much so it could then probably still get significant payload with reusability systems added on. So if it turned out that for this first stage use of an intermediate-sized launcher the mass ratio is that good, it would be a good test of the SSTO principle to launch the Starship without the upper stage to see how much payload it could carry as an SSTO. Bob Clark

The weight estimate is not new. It was the original weight estimate. How easy it would be is relative to taking another approach. The triple-core Falcon Heavy only cost 50% more in development cost than the Falcon 9 at triple the payload. In contrast, based on size, the 3 times larger SuperHeavy booster would cost 3 times more in development cost. Bob Clark

The increase in dry mass after the announcement of the addition of the movable wings was much discussed on the NasaSpaceflight.com forum. Prior to the new wings being added, Elon had said their specialty high-strength stainless steel version would require the same weight as the carbon fiber version. I agree it is puzzling why the increase would be that much. But there is precedent in movable wings being heavier as the example of the swing wing F-14 Tomcat shows. Bob Clark

No, I wasn’t referring to that larger diameter version, I think called the “Interplanetary Transport System”(ITS). I’m referring to essentially the current version but before the addition of the movable wings. It’s been much talked about on space oriented forums that the addition of the movable wings increased the Starship dry mass by approx. 50%, from 85 tons to 120 tons. I’m suggesting go back to the high mass ratio version. Then there are various options to get lightweight wings, landing gear, and thermal protection that would only subtract a proportionally small amount from the payload for reusability. Bob Clark

I’d like to get some feedback on the calculations here: Starhopper+Starship as a heavy-lift launcher. Triple-cored Starship for super-heavy lift. 2nd UPDATE, 9/2/2019: Starhopper as a lunar lander. https://exoscientist.blogspot.com/2019/07/starhopperstarship-as-heavy-lift.html Note: for the rest of this post, as well as in the blog post, I used the term “Starship” for its familiarity but I’m referring to the tanker version, not the version with the passenger quarters. The blog post argues that SpaceX should return to the originally planned high mass ratio version of the Starship, at ca. 25 to 1 rather than the current 10 to 1. The current mass ratio is quite poor for a dense-propellant rocket. SpaceX has been aiming to advance the state of rocketry, not go backwards. Then the Starship (the tanker version without passenger quarters) now used as a first stage plus a likewise weight-optimized Starhopper-sized upper stage could be a 100 ton class launcher in expendable mode. Note this would mean you would have a fully orbital-class launcher without the expense and extra time required developing the SuperHeavy booster. Then you would use triple cores a la the Falcon Heavy to get a 300 ton launcher. This approach would have multiple advantages. The biggest advantage is not needing the huge SuperHeavy at all. Because the Superheavy is three times the size of the Starship we can estimate its development cost as three times that of the Starship. In contrast, based on the Falcon Heavy experience, developing the triple-cored version would only cost 50% more than developing the Starship itself. The individual production cost would also be less, needing 1/3rd fewer engines. There is also the time element. Because of the Starhoppers small size and the fact it was already largely developed, aside from the required weight-optimizing, it could be produced along side the Starship at proportionally low cost. This is important because a 100 ton class launcher is commonly taken as the size-needed for a manned lunar mission. Then we could have a manned lunar mission mounted by next year in 2021 when Starship is expected to start flying. Another advantage is more controversial: the Starship, i.e., tanker version, at a 25 to 1 mass ratio and high Isp methane engines could be SSTO at significant payload. This would also be true for the weight-optimized Starhopper. This would go a long way to making manned-spaceflight routine since these smaller, simpler SSTO versions would be much more affordable and simple to operate and you could have independent companies and even private owners flying their own versions, both for point-to-point transport and for flights to LEO. The rocket equation shows the SSTO capability for the expendable mode. Here’s an argument that the extra weight needed for reusability, using weight optimized systems, would still allow significant payload as a SSTO: Short, stubby wings have been proven viable for return from space, so the large, heavy wings like the Space Shuttle are not required: The weight of the wings for the X-37 have not been revealed, however we can get an estimate from another vehicle the Skylon: For the Skylon the wing weight was only 2% of the landed weight. This is 2% of the full gross weight because it used a horizontal liftoff. But since the Starship will be using a vertical liftoff and non-lifting trajectory, the wings only have to support the weight of the vehicle on return, so that 2% will be calculated on just the dry weight. The landing gear weight can be taken as only 3%, or perhaps only 1.5%, of the dry weight: https://yarchive.net/space/launchers/landing_gear_weight.html Finally, the thermal protection as SpaceX’s PICA-X might only add on additional 8% of the dry weight. So these extra systems required for reusability will only add a proportionally small amount to the dry mass, so subtract only a proportionally small amount from the payload. Bob Clark

Isn’t it presumptuous to say they won when several companies have been granted contracts to develop a lander with the final decision to come later? Bob Clark

We can always be optimistic. There is no law that Boeing has to build the EUS, especially since it already has the contract for the SLS core stage. Perhaps we can start a letter writing campaign to Bridenstine to open up a competition for the EUS, like they did for the lander, with preference for the cryogenic upper stages already being planned. Bob Clark

An EUS can also mean a smaller lander that you want to send to the Moon with a single launch of the SLS. The Boeing EUS is unaffordable. But there will be several cryogenic upper stages that will be available as early as 2021 that could be used for a much cheaper EUS: Bob Clark

I speculated on a HLV using the Russian RD-171 engines here: https://exoscientist.blogspot.com/2012/05/low-cost-hlv.html Bob Clark

Thanks. I think the equations we discussed will be able to give the ideally optimized engine nozzle Isp’s. What needs to be worked on now are the trajectory equations including the gravity turn and the air drag equations. Any suggestions? Bob Clark

Thanks for doing the calculation. I assume you take into account that with fixed nozzles there is also a backpressure term given at the end of the equation here: F = q × Ve + (Pe - Pa) × Ae Also, as you know sea level engines are somewhat overexpanded at sea level because they also want to get good performance at vacuum. So the exit pressure for the fixed nozzle engine isn't really 1 atm at sea level. For the SSME's I think it's at around 1/3rd atm. For the Vulcain engines versions 1 and 2 the overexpansion is even worse. I think around at 1/4th atm for the Vulcain 1 and 1/5th atm for the Vulcain 2. You can get an idea about the degree of overexpansion for the Vulcain 1 at sea level from this graphic: At sea level the ideal Isp is significantly better than what the Vulcain 1 gives. I did a rough estimate and found for the currently-used Vulcain 2 because of its even higher level of overexpansion, its sea level thrust could be increased ca. 30%(!) by given it ideal expansion at sea level. This is important because higher lightoff thrust can reduce gravity drag. The increase for most engines wouldn't be this great though because most aren't this greatly overexpanded. For some other engines I tried I estimated the increase was less than 10% better sea level thrust by giving it ideal sea level expansion. For the parameters that go into the equation for the exhaust velocity, combustion temperature, molecular weight of combustion products, specific heat, etc., I've used the shareware program Rocket Propulsion Analysis, http://propulsion-analysis.com/index.htm. Bob Clark

Assuming we can use the equation for the exhaust velocity ve dependent on ambient pressure, we still need to calculate the flight path to orbit. For simplicity sake we can use the “gravity-turn” trajectory: Gravity turn A gravity turn or zero-lift turn is a maneuver used in launching a spacecraft into, or descending from, an orbit around a celestial body such as a planet or a moon. It is a trajectory optimization that uses gravity to steer the vehicle onto its desired trajectory. It offers two main advantages over a trajectory controlled solely through the vehicle's own thrust. First, the thrust is not used to change the spacecraft's direction, so more of it is used to accelerate the vehicle into orbit. Second, and more importantly, during the initial ascent phase the vehicle can maintain low or even zero angle of attack. This minimizes transverse aerodynamic stress on the launch vehicle, allowing for a lighter launch vehicle.[1][2] https://en.m.wikipedia.org/wiki/Gravity_turn#Launch Doing a google search turned up several references on calculating gravity turn trajectories. For this first level analysis I’m looking for some easily implemented ones if anyone knows of any. Also, some kerbal RealSolarSystem mod simulations have been done using aerospike nozzles for SSTO’s. Aerospike nozzles are the closest we have to ideal adaptive nozzles. Anyone want to give it a try with aerospike engines replacing the engines on the F9, Delta IV, Atlas V or other first stage boosters? Bob Clark

Dragon01 mentioned the equation for Isp was linear. It’s not really, but this reminded me of something I’m puzzled about. Fixed nozzles on a sea level engine are a compromise. They are overexpanded for sea level operation so they can get good Isp in vacuum.This should mean they get optimal Isp at some intermediate altitude, not at sea level and not in vacuum. But actually in graphics of engines they show the Isp either constant or increasing towards vacuum conditions. The graphic would be expected instead to look like this: Taken from this page: http://www.braeunig.us/space/sup1.htm But using the first image above we might be able to model approximately the Isp according to altitude by two straight lines, both for the fixed nozzle case and for the adaptive nozzle case. For the fixed nozzle case it would be an inclined straight-line to the altitude of 15,000m, then switching to a constant, i.e., flat-line thereafter at the vacuum Isp value. For the adaptive nozzle case, it would be two inclined straight lines. The first would be steeper than the second with the transition at around 15,000m to 20,000m. Bob Clark