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Monday, January 16, 2023

Can we adapt the Antarctica IceCube Neutrino Observatory drilling technique to explore the subice oceans of Europa?

 Copyright 2023 Robert Clark


 The IceCube Neutrino Observatory was able to drill down 2.5 kilometers rapidly into the ice in Antarctica using a relativly simple method using heated, pressurized water. Can we adapt this to reach the subice oceans on Europa?


 Here is a published description of the drilling method:

IceCube Enhanced Hot Water Drill functional description.
Published online by Cambridge University Press:  26 July 2017
T. Benson, J. Cherwinka, M. D
https://www.cambridge.org/core/journals/annals-of-glaciology/article/icecube-enhanced-hot-water-drill-functional-description/F2A93B250A9FDB4E8A94101787BA3A68

 The specifications:


 The total power for drilling a 60 cm wide hole was 5 MW. The great majority of this was the thermal power of 4.7 MW for melting the ice. 

 For just sending a small probe beneath the ice we'll make the hole width smaller by a factor of 10 to 6 cm, 2.4 inches. A probe of greater capabilities will be obtained by simply making it longer, while keeping it the same width.

 The required power will scale by the cross-sectional area to 50,000 watts, 50 kW. 

 In this blog post I noted the Falcon Heavy can send a 1 ton spacecraft to the surface of Europa using currently existing in-space stages:


 The Europa Clipper mission currently planned for launch in 2024 on the Falcon Heavy is an orbiter mission though; it will not have a lander. However, the cost of the Falcon Heavy is only ca. $100 million, so a second launcher for the lander mission can easily be purchased. By the way, an interesting and much appreciated phenomenon I've observed about NASA science missions is the time between announcement and actual mission launch seems to much more rapid than before. So it's likely the lander mission can be launched also within the ca. 2024 time frame.

 There have been several proposals to drill through the Europan ice. But they frequently suggest doing it over long time frames, several months. For instance, the proposal discussed at about the 8:50 point in this video would take up to a year:



 The IceCube drill though took only 30 hours to go through 2.5 kilometers of ice in Antarctica. For a 15 kilometer thick ice shell on Europa that would be about 7.5 days. However, because of the fractured surface of Europa the ice shell is believed to be much thinner in some place, perhaps as low as 5 km thick. In that case, the drill may only take 2.5 days to penetrate the ice.

 The REASON ice-penetrating radar on the Europa Clipper orbiter can be used to first determine thinnest locations on the ice for the lander to land at.

Supplying the Energy.

 But first how to get 50 kW power at Europa? We have a limit of about 1 km for the lander spacecraft payload to be launched with in-space stages on the Falcon Heavy. First, space fission reactors. Here's a list of space fission reactors that have flown in space:

Name and modelUsed on (# of RTGs per user)Maximum outputRadio-
isotope
Max fuel
used (kg)
Mass (kg)Power/mass (Electrical W/kg)
Electrical (W)Heat (W)
MMRTGMSL/Curiosity rover and Perseverance/Mars 2020 roverc. 110c. 2000238Puc. 4<452.4
GPHS-RTGCassini (3)New Horizons (1)Galileo (2)Ulysses (1)3004400238Pu7.855.9–57.8[48]5.2–5.4
MHW-RTGLES-8/9Voyager 1 (3)Voyager 2 (3)160[48]2400[49]238Puc. 4.537.7[48]4.2
SNAP-3BTransit-4A (1)2.7[48]52.5238Pu?2.1[48]1.3
SNAP-9ATransit 5BN1/2 (1)25[48]525[49]238Puc. 112.3[48]2.0
SNAP-19Nimbus-3 (2), Pioneer 10 (4)Pioneer 11 (4)40.3[48]525238Puc. 113.6[48]2.9
modified SNAP-19Viking 1 (2), Viking 2 (2)42.7[48]525238Puc. 115.2[48]2.8
SNAP-27Apollo 12–17 ALSEP (1)731,480238Pu[50]3.8203.65
(fission reactor) Buk (BES-5)**US-As (1)3000100,000highly enriched 235U3010003.0
(fission reactor) SNAP-10A***SNAP-10A (1)600[51]30,000highly enriched 235U4311.4
ASRG****prototype design (not launched), Discovery Programc. 140 (2x70)c. 500238Pu1344.1

 Note to get to 50 kW would require several hundred kg just for the power system alone, leaving little for the rest of the spacecraft operations. One might also not like to leave radioisotopes on a world that may have life.

 Solar power? Normally the high power requirements would make this a non-starter for the far away Jupiter system with its reduced solar insolation. But quite key here is the majority of the power needs to be thermal. The inefficiencies of space solar power come from when it has to be converted to electrical power. Jupiter gets about 50 watts per square meter of solar power, well less than the 1,200 W per square meter at Earth. Then you would need 1,000 square meters solar collecting area, about 32m x 32m. But you would not send this to solar electric cells. You would collect it either by a parabolic mirror or Fresnel lens to concentrate it to achieve a temperature of up to 3,000°C. In the low gravity of Europa this would need minimal support structures and might weigh only a few tens of kilos total weight. For instance the Sunjammer solar sail only weighed 32 kg at a 1,200 square meter area

 Burning fuel? The disadvantage here is that higher fuel load would need to be carried to Europa for a longer time the drill had to be run. To estimate the power using a rocket engine to deliver the thermal power. it's at 1/2 times the square of the exhaust velocity per fuel flow rate. Or equivalently, the amount of energy per kilogram propellant is 1/2 times the square of the exhaust velocity. Given a (very) large expansion ratio the exhaust velocity of a hydrolox engine can be ca. 4,800 m/s. So the energy per kilo of propellant would be 11.5E6 J/kg. If we ran our 50 kW thermal power system for 7.5 days for a 15 km thick ice shell, it would take 50,000*7.5*24*3,600 = 32.4E9. This would take 2,800 kg of propellant, far too much.

Estimates of power requirement for Europa. 

 The above was assuming the IceCube drilling method could be adapted to Europa, specifically the energy requirements. This page actually does the calculation from first principles for Europa:

Calculating the Required Energy of Drilling into Europa.
July 9, 2020 by Engineer Calcs

 It discusses three methods: 1.)breaking up the ice and simply carrying the broken up ice to the surface, 2.)bringing the  -160°C ice on Europa only up to a easier to achieve sublimation temperature of -75°C rather than to the melting point of 0°C, and allowing the ice to sublimate to a gas in the vacuum of Europas surface, and 3.)bringing the ice up to the 0°C melting point and using analogous methods as to the IceCube experiment.

 The first method requires surprisingly low amounts of energy. For his calculation, the author takes the width as 1-meter and the depth 25 kilometers. The enegy requirement is simply calculated by the potential energy of raising that much mass of ice the required height in the low gravity of Europa. He calculates 5.92E11 Joules. However, our case will be smaller in using only a 6 cm wide drill hole and assuming at most 15 km height. Also, in actuality not the entire column needs to be raised the full height. The ice near the top for example needs to be raised hardly at all. The height and the energy needed to raise is a linear function of the depth of that portion of the ice. Then the energy can be calculated by using just the middle point of the height, i.e., only 7.5 km for our 15 km drill hole. Then the energy is:

5.92E11*(6/100)^2*(7.5/25) = 6.4E8 Joules.

 Using our 50,000 watt solar generator, that would only take 3.4 hours to generate the required energy(!) One question though, the energy needed in this case would be mechanical rather than thermal in mechanically raising the ice to the surface. What would the conversion efficiency be? Because of the quite high temperature of 3,000°C, by the Carnot Law the conversion efficiency could be quite high, certainly over 90%.

 Because the energy requirements for the mechanical lifting method are comparatively low we might even allow the fuel burn method of supplying the power rather than using solar thermal. Our needed total energy now is 6.4E8 joules. Using the calculated energy we used for a hydrolox engine of 11.5E6 J/kg of propellant, that would only require 6.4E8/11.5E6 = 57 kg.

 However, a problem with hydrolox with long space missions is boiloff. There is research in reducing it but whether it can be reduced enough for a misson of 3 to 4 year travel time through space is uncertain. So we'll estimate it using storable hypergolic propellants, same as used for the propulsion stages at Jupiter. In vacuum these can have 3,200 m/s exhaust velocity. This can give (1/2)*3,200^2 = 5.12E6 J/kg energy. Then it would take 125 kg of the storable propellant. This would still be doable for a 1,000 kg spacecraft payload. 

 The author of the "Calculating the Required Energy of Drilling into Europa" page does not prefer the mechanical lifting method, preferring to use the melting ice method eventhough it would take 2 orders of magnitude more power. This is because with a mechanical lift it may have a tendency of breakdwon operating in such low temperatures over long time.

 However, as the previous calculation using the solar thermal generator above showed it might be it would only take on the matter of hours. This would certainly reduce the likelihood of breakdown.

 Note also using the fuel-burn method of generating power, the moving exhaust itself could be used to raise the ice to the surface, thus elimating the issue of mechanical breakdown of the lifting method.

 There is another issue though for raising the ice to the surface without expending extra energy to melt it. Ice is brittle so likely would require low energy compared to the energy needed to raise to the surface to break it up. This is the case on Earth. However for the extreme level of ice compaction of Europa the ice would be denser so might be harder to break up. Experiments would need to be done to determine this.

  That discussion was for the 1st method of drilling down of mechanically raising the ice to the surface.. For the second method of only raising the ice to sublimatioin temperture of -75°C at first glance that might appear to be preferable. However, the energy of sublimation, the energy required to sublimate from ice to gas is so high, about 10 times higher than that just to melt it, it winds up taking more total energy than raising the temperature of the ice to the melting point of 0°C.

 So we'll consider the 3rd method of melting the ice and then using analogous methods as in the IceCube experiment. The calculation of the energy in this case is 1.2E13 but this was for a 1-meter wide drill hole and for a 25 km drill depth. So for our 6 cm wide drill hole down only 15 km the energy required would be: 1.2E13*(6/100)^2*(15/25) = 26E9. Using 50 kW solar thermal power, this would only take about 6 days. 

 Actually, in his calculation the author of the "Calculating the Required Energy of Drilling into Europa" page used a constant temperature of  -160°C for the ice that had to be raised to 0°C for melting. But since at the water level below the ice, the temperature will already be at about 0°C, perhaps a little less for salt freezing point depression, the temperature will gradually decrease linearly from 0°C to -160°C at the surface. So the calculation needed for the energy needed for the temperature change can be done by using the average of -80°C. This cuts the energy required for raising the ice to the melting point by about half. However, the energy for the phase change from ice to liquid will still be the same, resulting in the total energy being in the range of 75% of the estimated energy above. 

 Also, importantly, in comparison to the IceCube experiment drilling method, which took place in Antarctica at -50°C, an average temperate of -80°C on Europa is only moderately far from -50°C, compared to the -160°C temperature at the surface of Europa, which gives us confidence the energy requirements would be in line with those of the IceCube experiment.

 The author of that page prefers this melting method and we now have a good point of comparison to the IceCube experiment to have good confidence that it can actually work. As I mentioned before a great point in its favor is the quite short drilling time of only a fews days compared to the months to a year estimated for other proposals.

 Still, I am intrigued by the low energy requirements of  just mechanically raising of ice to the surface without melting, especially if using the exhaust from fuel-burn method of raising the ice to the surface, since this could shorten the drill time down even further to hours instead of days.

   Robert Clark 
 

  

Monday, January 9, 2023

Aerospike in 3D exhaust injection. UPDATED, 1/10/2023: Extension to single nozzles.

Copyright 2023 Robert Clark

(patents pending)


 The aerospike is a method of altitude compensation for a rocket nozzle that provides the optimal expansion for the ambient air pressure all the way to orbit. Two ways of doing this is one, to use a toroidal combustion chamber and exhaust the flow down the sides of the spike.

 Another way is to use multiple small combustion chambers that each separately exhaust down the sides of the spike. 


 In both cases the exhaust flow is initiated at only the top of the spike. However, a problem is converting a cylindrical combustion chamber to a toroidal one is an expensive process. 

 Another problem would be for the case of a Falcon 9 or a SuperHeavy where the entire base area is covered by the multiple engine bells there would be no room to place a spike below it. 




 But a way it could be done would be to arrange the various engines in three dimensions down the sides of the spike:


UPDATE, 1/10/2023           

 The above was for a multi-engine stage. But there is a way to emulate this for a single engine to produce a 3D level exhaust flow as for the multi-engine case. 

 Imagine concentric rings attached to the bottom of the exhaust nozzle  of the single engine



The exhaust would flow between the separate rings. Now imagine each ring extended down with the length of each ring extending down more as you proceed further inward. 


 The exhaust flow would follow the exterior of an aerospike shape as before.

   Robert Clark

Thursday, November 10, 2022

A low cost, lightweight lunar lander.

 Copyright 2022 Robert Clark


 In the blog post Possibilities for a single launch architecture of the Artemis missions I discused that a single launch architecture is possible for the SLS rocket if there is a lightweight lunar lander. Such is possible using currently existing space stages. Firstly, a lunar crew module of ca. 2 ton mass is possible based on Orbital Sciences Cygnus capsule, discussed in Budget Moon flights: lightweight crew capsule. The Cygnus is actually build in Italy by Thales Alenia Space. As it is already built, the additional modications for added life support would be comparatively low cost.

 Note Thales Alenia Space is already adding life support sysmtems to a larger version of the Cygnus for the lunar Gateway. Then a lower cost version would simply use the smaller Cygnus itself, at a ~2 ton dry mass for a short term stay on the lunar surface.

 As for the propulsion system, the earlier Ariane 5 EPS storable propellant upper stage prior to the current cryogenic upper stage could be used for the purpose: it had a 1.275 ton dry mass and 9.750 ton propellant mass, for a 11 ton gross mass. Then the crew module and propulsion stage would mass 13 tons.


 An advantage over the SpaceX Starship lunar lander plan is that it is only 3 meters high, the same height as for the Apollo lunar lander descent stage, making it easy for the astronauts to climb down to the lunar surface, compared to the 25 meter height for the Starship.

 Calculation.

The delta-v to the lunar surface from low lunar orbit is 1,870 m/s:


 The Aestus engine on the stage has a Isp of 324s. Then the delta-v it could achieve carrying a 2 ton crew module would be:

324*9.81Ln(1 + 9.75/(1.25 + 2)) = 4,400 m/s. Then it could work as single stage to go down to the lunar surface from low lunar orbit and back again.

 It is notable a stage derived from the Space Shuttle OMS pods would also have this capability.


 The specifications for the OMS pods are given here:

SHUTTLE PERFORMANCE ENHANCEMENTS USING AN OMS PAYLOAD BAY KIT 1991

The middle size version has a dry mass of 3,955 lbs, 1,800 kg, and propellant load of 25,064 lbs, 11,400 kg. It has an Isp of 316 s. Then with a 2 ton crew module it would have a delta-v of 4,300 m/s:

316*9.81Ln(1 + 11.4/(1.8 +2)) = 4,300 m/s, sufficient for single-stage lunar lander.

  Robert Clark


Saturday, November 5, 2022

Why does the Boeing Exploration Upper Stage(EUS) cost so much?

 Copyright 2022 Robert Clark


 The Boeing Exploration Upper Stage(EUS) is only about 1/10th the size of the SLS core stage. Why couldn't this upper stage be built in the same fashion as the core?

 The SLS core like the shuttle External Tank before it is built by welding together barrel sections:


 Then the upper stage could have been formed in the same fashion simply by using fewer/shorter barrells. Indeed, cost could have cut even further by constructing the upper stage at the same time as the core.

 Being only ~1/10th the size of the core it likely would have fit in the same building during the construction:


 Instead of this simple approach, Boeing chose a more complicated construction that resulted in higher development costs:


  In fairness to Boeing it should be said that they most likely wanted to save on dry mass on the stage. Unlike for lower stages like the core, extra mass on the upper stage results in direct reduction in that mass from the payload.

 However, to accomplish the same payload with a less mass efficient stage they could have just increased the stage propellant size. For instance the Delta IV first stage has a 200 ton propellant load at a 26 dry mass. You would not need the full thrust and high mass of the sea level engine the RS-68 used on the stage for this upper stage use. Smaller thrust and weight engines such as 4 RL-10's or a single J-2X would suffice.

 The mass efficiency for the Delta IV stage without the 6 ton RS-68 engine is also pretty good at a 200 ton propellant load for the 20 ton structural mass sans engine. Note the Delta IV first stage achieves this good mass efficiency without using advanced mass saving methods such as common bulkhead, balloon tanks, or specialty high strength stainless steels as used on the Centaur upper stages.

BOEING EUS COSTS.

 The total development and production cost for the Boeing EUS will probably be $10 billion:


More far-out facts: The EUS will probably cost $10 billion and (it's being built by Boeing, recall) won't be completed for at a minimum of five years. Do you think a fully expendable $2.5 billion super heavy lift rocket will have any relevance at that time?

https://twitter.com/SciGuySpace/status/1446478856840433669?s=20&t=LILNIi_t_RepDRJMMTX49Q 

And the per stage cost likely ~$880 million each:

The story also uses a simple cost estimator model to put a per-unit price on Boeing's Exploration Upper Stage. The result? $880 million. That's just the upper stage, and doesn't include the core stage, SRBs, integration, ground systems, etc. etc.

 Note the entire SLS development was in the range of $20 billion, with an expected per lauch cost of $2 billion. Then this EUS stage at ~1/10th the size of the SLS core and without the extra expense of the SLS side boosters or current interim SLS upper stage would amount to half the cost of the entire current SLS.


ENGINE COSTS.

 While the RS-68 engine(with a vacuum optimized nozzle extension) would be over powered for the upper stage use, it would be significantly cheaper than using 4 RL-10's for example. The RL-10's are estimated to cost $25 million for a total of $100 million for 4. The RS-68 in contrast is estimated to only cost $10 to $20 million. For the J-2X engine, being an engine only developed but not yet used, there are no estimates on its cost, but it's very likely to cost much less than the $100 million for the 4 RL-10's. Another possibility might be the Vulcain engine used on the Ariane 5. It's estimated cost is only $10 million.

 As an estimate on how much the stage aside from the engines might cost if built in the same fashion as the SLS core and shuttle external tank, look at the costs for the shuttle ET built by Lockheed. They were given a contract for 18 External Tanks for a total $3 billion, about $166 million per tank. Note too an SLS upper stage at ~200 ton propellant load would be less than 1/3rd the size of the External Tank. Given this much lower cost Lockheed should be asked how much would be a stage built in the same fashion as the shuttle ET at ~ 200 ton propellant load.


  Robert Clark