SpaceX should withdraw its application for the Starship as an Artemis lunar lander, Page 3: Starship has radically reduced capability than promised.

Copyright 2024 Robert Clark

 Elon Musk presented an update on the plans for the Starship post the third Starship test flight:

Elon Musk Starship Presentation: IFT-4 Master Plan, Starship V2 & V3, Raptor V3, Mars, IFT-3 & More.

 About 31 minutes in Elon suggests the current version V1 would be capable of 40 to 50 tons to orbit. This is bad because SpaceX sold NASA on the idea the Starship HLS could serve as an Artemis lander based on 150 tons to orbit reusable and “10ish” refueling flights. If the capability is max 50 tons, then it would take “30ish” refueling flights.

 If they intend to use version V2 then this is bad because it would require further qualification flights for the larger version and more importantly further qualification of the more powerful Raptor 3 engine needed.

 This last is doubly bad because I’d be willing to bet dollars to donuts that they never informed NASA that the current version couldn’t do it and further development would be required for the larger version.

  SpaceX needs a true Chief Engineer. Elon once said that early on when there were still doubts about its viability, they tried to recruit a Chief Engineer for SpaceX but no one good was willing to come. So Elon designated himself Chief Engineer. It is not a role Elon is well suited for. A good Chief Engineer should be scrupulously forthright. He would not refer to the little 5 or 10 second static burns SpaceX does for the SuperHeavy or Starship as "full duration".

 A true Chief Engineer would be aware that "full duration" in the industry is short for "full mission duration". These static fires in the industry are conducted at the full length and the full thrust of an actual flight and are meant to give confidence to potential customers that the engines can perform as expected for the promised capabilities of the launchers.

 However, SpaceX in using the term "full duration" for these little few-second burns, doesn't even tell the public, or its major customer NASA for which they have a billion-dollar contract, if these little burns are even conducted at full thrust.

 This has had majorly negative consequences. The FAA had great concerns in the Raptor reliability after the first test flight. In the "corrective actions" they required of SpaceX prior to a second Starship test flight, at the top of the list was correcting the tendency of the Raptor of leaking fuel and catching on fire while in flight.

 I have argued multiple, independent lines of evidence suggest SpaceX intentionally reduced the throttle of the Raptors on the booster on the second test flight, IFT-2, to improve reliability of the engines:

Did SpaceX throttle down the booster engines on the IFT-2 test launch to prevent engine failures?https://exoscientist.blogspot.com/2023/12/did-spacex-throttle-down-booster.html

 Running an engine at reduced throttle reduces the pressure levels within the engine, high pressure being a major cause of engine fuel leaks. The Starship upper stage though was run at near full throttle on IFT-2, perhaps because performance would be reduced too much if it also was run at reduced throttle.

 The result was the booster engines worked fine, at least on ascent, while the Starship exploded on ascent on IFT-2. SpaceX has said the Starship RUD was due to an intentional LOX dump they performed to keep that flight as suborbital. However, many knowledgeable observers doubted the LOX dump alone would have caused a RUD. They argue due to the tendency of the Raptor to leak fuel, it's more likely that plus the LOX dump caused the RUD.

  For the third test flight, IFT-3, after reviewing both propellant burn rates and the acceleration profile of the flight, I'm suggesting SpaceX learned their lesson from the second test flight, and this time both stages were run at reduced throttle on this flight. And this time both stages were able to complete the ascent stage of the flight successfully.

 However, this does reduce the payload capability of the launcher. Elon has acknowledged this radically reduced payload capability in his recent update. But it needs to be explained by SpaceX why the payload is so greatly reduced. If it is because the Raptor needs to be run at reduced thrust in order to be reliable then that is an extremely important thing to acknowledge, and to inform NASA on it, because the thrust levels of a rocket go into assessing what its actual capabilities are.

 There is another very important issue about Raptor reliability. Multiple times a Raptor has undergone a RUD doing a relight during prior testing of the Starship planned landing procedure. And on this last Superheavy/Starship test flight as well Raptors underwent a RUD during the booster landing procedure. The boostback back burn appeared to have occurred successfully. But there was venting gas after the bostback back burn suggesting there may have been a fuel leak here as well.

 Note for a successful reuse of the Starship and booster, successful relights have to occur both for boostback burns and landing burns. Then in none of the prior Starship landing tests nor of the Superheavy/Starship flight tests have any flights shown successful Raptor relights without leaking fuel and catching fire, and often undergoing a RUD. 

 SpaceX has called one test of the Starship landing test, SN15, successful because it managed to land without exploding. But it is important to note even in that test a Raptor leaked fuel and caught fire prior to landing. It's just on that test SpaceX managed to extinguish the fire before the ship exploded:

 Note that in the SpaceX plans for a reusable Starship it absolutely can not work if the Raptor can not be made to relight reliably. SpaceX in not publicly providing full mission duration, full thrust testing information on the Raptors have not shown this also for relights of the Raptor.

 That is why it is so important for a launch company to publicly provide details on full mission duration, full thrust level static engine testing.

 SpaceX needs a true Chief Engineer to provide such details in a forthright manner.

    Robert Clark

SpaceX should explore a weight-optimized, expendable Starship upper stage.

 Copyright 2024 Robert Clark

 To me it’s just stunning SpaceX is ignoring that an expendable Starship could be done for 40 ton dry mass, choosing instead the current 120 tons for the reusable version: 

Elon Musk

@elonmusk

Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible.

9:14 PM · Mar 29, 2019

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https://twitter.com/elonmusk/status/1111798912141017089

 Keep in mind that every kilo of extra mass in an upper stage subtracts directly from the payload possible. Then that 80 tons difference in the dry mass between the reusable and expendable versions is a huge difference. 

 Now,  note because of size, that, just like with the Falcon 9, the 1st stage is 2/3rd of the cost. So for ~$90 million total for the SuperHeavy/StarShip, the SuperHeavy is $60 million of that. But as the Falcon 9 shows it is much easier to get reusable 1st stage. So assume with reuse of SuperHeavy, its cost, is now, say, $5 million per launch. Now it’s a $35 million total cost for the partially reusable SuperHeavy/StarShip. BUT now because of the radically reduced upper stage dry mass, we have ca. 300 tons payload this version!(Assume SuperHeavy lands down range if you wish to maintain the high payload.) But this is about the same cost per kilo as fully reusable 100 to 150 ton payload fully reusable version at $10 million per flight cost.

 Then the question is how realistic is it the Starship could have 40 ton dry mass as an expendable? I think it is quite realistic. 

 Consider the original Atlas rocket first used to send John Glenn to orbit:

SLV-3 Atlas / Agena B.

Family: Atlas. Country: USA. Status: Hardware. Department of Defence Designation: SLV-3.

Standardized Atlas booster with Agena B upper stage.

Specifications

Payload: 600 kg. to a: 19,500 x 103,000 km orbit at 77.5 deg

inclination trajectory.

Stage Number: 0. 1 x Atlas MA-3 Gross Mass: 3,174 kg. Empty Mass:

3,174 kg. Thrust (vac): 167,740 kgf. Isp: 290 sec. Burn time: 120 sec.

Isp(sl): 256 sec. Diameter: 4.9 m. Span: 4.9 m. Length: 0.0 m.

Propellants: Lox/Kerosene No Engines: 2. LR-89-5

Stage Number: 1. 1 x Atlas Agena SLV-3 Gross Mass: 117,026 kg.

Empty Mass: 2,326 kg. Thrust (vac): 39,400 kgf. Isp: 316 sec. Burn time: 265 sec. Isp(sl): 220 sec. Diameter: 3.1 m. Span: 4.9 m. Length:

20.7 m. Propellants: Lox/Kerosene No Engines: 1. LR-105-5

Stage Number: 2. 1 x Agena B Gross Mass: 7,167 kg. Empty Mass: 867 kg. Thrust (vac): 7,257 kgf. Isp: 285 sec. Burn time: 240 sec. Isp(sl): 0 sec. Diameter: 1.5 m. Span: 1.5 m. Length: 7.1 m. Propellants: Nitric

acid/UDMH No Engines: 1. Bell 8081

http://www.friends-partners.org/partners/mwade/lvs/slvgenab.htm

 The Atlas had an unusual design however. It dropped its main lift-off engine at altitude and continued on with what was called the “sustainer” engine. This engine due to much of the propellant mass being burned off had much lower thrust, and so much reduced required engine weight. Then looking at the specifications of this stage, note it had nearly a 50 to 1 mass ratio(!)

 The comparison of this sustainer stage to the 3-engine Starship upper stage is appropriate since an upper stage typically doesn’t need to have the thrust of a stage needing to lift off from the ground. Weight growth of the Starship now at 120 tons dry mass required adding 3 additional engines, to now have 6 engines.

 However, a key reason why the Atlas was able to achieve such a high mass ratio was that it used what was called “balloon-tank” design. This was a design that used pressurization to maintain its structure even on the ground. It would actually collapse under its own weight when not pressurized.

 However,  methanolox is at about 80% of the density of kerolox. So a corresponding methanolox version would be at 40 to 1 mass-ratio, better than the 30 to 1 mass ratio Elon suggested. But its not likely SpaceX would want to deal with the operational difficulties of having a stage be continually pressurized even when on the ground, unfueled, especially for a stage intended to have high launch rates.

 So I’ll look at another stage, the S-II hydrolox 2nd stage of the Saturn V rocket. The Saturn V launcher of the Apollo program was remarkable in the lightweight features of its upper stages, the S-II and the S-IVB. This page gives a list of the fueled weights and empty weights of the Saturn V stages:

Ground Ignition Weights

http://history.nasa.gov/SP-4029/Apollo_18-19_Ground_Ignition_Weights.htm

 The later versions of Apollo had improved weight optimization. We'll use the specifications for Apollo 14. The "Ground Ignition Weights" page gives the Apollo 14 S-II dry weight as 78,120 lbs., 35,510 kg, and gross weight as 1,075,887 lbs., 489,040 kg, for a propellant mass of 997,767 lbs., 453,530 kg, resulting in a mass ratio of 13.77 to 1. 

 Now, methanolox is 2.5 times greater density than hydrolox. Then the corresponding mass ratio for methanolox would be at 33 to 1. This comparison is particularly apt because the mass in the same size tanks would be approx. at the 1,200 propellant mass of the Starship.

 So Starship could reach ca. 30 to 1 mass ratio when using the weight optimizing methods used during the Apollo program.

 But if the price per kilo of this partially reusable version would be at about what the current version is what is the advantage? One advantage is as mentioned is you would not have the difficulty of making the upper stage reusable, no problematical heat shield tiles.

 There is another advantage not as concrete, but in my mind just as important if not more so. In my opinion the approach SpaceX is taking with the SuperHeavy/Starship is ill-conceived. It is based on the idea the SuperHeavy/Starship should be the be-all-end-all for ALL of spaceflight.

 But if you look at transport methods throughout history even going back to the horse-drawn era transports always came in different sizes. A comparison to the air traffic is most instructive. It turns our the largest air transports the jumbo-jet size aircraft actually make up a tiny percentage of air traffic. The great bulk of air traffic is carried by smaller aircraft.

 And even looking at SpaceX’s own Falcon Heavy demonstrates this. The per kilo cost is less than that of the Falcon 9. But the number of Falcon Heavy flights is tiny compared to the number of Falcon 9 flights.

 The fixation on the reusable Starship as the be-all-end-all for all spaceflight also leads to the poorly-conceived notion that a Mars or Moon mission must be carried out by multiple refuelings of the reusable Starship. The number of refueling flights for the Artemis lunar missions might be 8 to 16 flights.

But it is a basic principle of orbital mechanics that high delta-v missions such as to the Moon or Mars are more efficiently carried out by using additional stages. Simply by giving the SuperHeavy/Starship an additional 3rd stage, flights to both the Moon and to Mars could be carried out in a single launch.

 An expendable Starship would mean it being regarded as just another stage. And a 3rd stage could be set atop it as needed, such as for high delta-v missions. 

 As another illustration of the fact this approach to the SuperHeavy/Starship is ill-conceived, the payload of the SH/ST to GEO is nearly zero because that Starship dry mass is so high. This is the most lucrative satellite market, but a single SH/ST launch could not service that market. In order to just launch satellites to GEO the SH/ST would have to do multiple refuelings just to launch a satellite to GEO, just like when it had to launch manned interplanetary missions. This is an odd state of affairs for a rocket simply to launch satellites to GEO.

 Or of course it could utilize a 3rd stage. But if you are going to use a third stage then, why not just use it also for the manned interplanetary missions that would allow you to do such missions in a single flight?

 Robert Clark

Altitude compensation is more efficient than staged-combustion engines.

 Copyright 2024 Robert Clark

 Staged combustion engines such as the Russian RD-180 and American SSME are regarded as the utmost in efficiency because they achieve high vacuum Isp while being able to achieve high thrust at sea level. They achieve this by operating at high chamber combustion pressure. 

SpaceX is developing the Raptor engine also as a staged combustion engine. However, a surprising fact is a medium performance, mid-level pressure and cheaper engine such as the Ariane 5’s Vulcain engine or Delta IV’s RS-68 can get higher performance than a staged combustion engine by using altitude compensation.

See the graphic of the Isp of the Vulcain engine with an altitude compensating nozzle:

 The vacuum Isp of the SSME is 452.3 seconds (4,436 m/s), and the sea level, 366 seconds (3,590 m/s). You see from the graphic with altitude compensation the Vulcain sea level Isp would be ca. 3,850 m/s. And already at ca. 20,000 m, its altitude compensating Isp would match that of the SSME. An thereafter the the Isp would exceed the maximum vacuum Isp of the SSME. Indeed its Isp could reach 4,850 m/s, and above

That the Isp with adaptive nozzles can be this high is supported by calculations for hydrogen/oxygen engines at ultra large expansion ratios. This report concludes at a 600 to 1 expansion ratio we can get ca. 480 s vacuum Isp:

ORBITAL TRANSFER VEHICLE (OTV) ENGINE STUDY, PHASE A - EXTENSION
CONTRACT NO. NAS8-32996

https://ntrs.nasa.gov/api/citations/19800015867/downloads/19800015867.pdf

 The method to get the altitude compensation does not have to be the aerospike nozzle. Better actually would be to add an altitude compensation nozzle extension to an existing engine such as the Vulcain or RS-68. Redesigning such an engine to use an annular combustion chamber for an aerospike nozzle would be expensive. Far cheaper would be to use an altitude compensating nozzle attachment to the already existing engine.

 

Such nozzle extensions already have been in existence for decades on upper stage engines, such as the extendable nozzles on for example the RL-10B2 engine. 

 But in actually the increase in efficiency would be higher for a first stage engine. For instance the vacuum Isp for the Vulcain or RS-68 could be increased from 432s or 412s to 480+ s and above.

 

 The nozzle extension is just a well-known, simple way to accomplish it but there may be simpler or more lightweight methods of accomplishing it. 

Aerospike in 3D exhaust injection. UPDATED, 1/10/2023: Extension to single nozzles.
https://exoscientist.blogspot.com/2023/01/aerospike-in-3d-exhaust-injection.html

SSME based SSTO’s. UPDATED, 6/28/2021 - Extension to the Delta IV Heavy.
https://exoscientist.blogspot.com/2021/06/ssme-based-sstos.html

ESA's Callisto reusability testbed as an *operational* TSTO and SSTO. UPDATE, 7/1/2019.
https://exoscientist.blogspot.com/2019/05/esas-callisto-reusability-testbed-as.html

Altitude compensation attachments for standard rocket engines, and applications, Page 6: space shuttle tiles and other ceramics for nozzles. UPDATED: 3/6/2018
https://exoscientist.blogspot.com/2017/12/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 5: metal foil expandable nozzles.
https://exoscientist.blogspot.com/2017/08/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 4: the double aerospike.
https://exoscientist.blogspot.com/2016/10/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 3: stretchable metal nozzles.
https://exoscientist.blogspot.com/2016/06/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications, Page 2: impulse pressurization methods.
https://exoscientist.blogspot.com/2016/01/altitude-compensation-attachments-for.html

Altitude compensation attachments for standard rocket engines, and applications.
https://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html

The Coming SSTO's.
https://exoscientist.blogspot.com/2012/05/coming-sstos.html

Altitude Compensation Improves Payload for All Launchers.
https://exoscientist.blogspot.com/2016/01/altitude-compensation-improves-payload.html

 

  Robert Clark

Could meteor impacts be the cause of the coronal heating problem?

 Copyright 2024 Robert Clark

 A puzzle in solar science that has existed for 150 years is the corona heating problem:

Why is the sun’s corona 200 times hotter than its surface?
The paradox has astronomers scratching their heads over magnetic waves, nanoflares, and the now-debunked element coronium.
BY BRILEY LEWIS | PUBLISHED APR 12, 2023 6:00 AM EDT
https://www.popsci.com/science/how-hot-is-the-suns-surface-corona/

 The Sun's surface is at about 10,000 F, 5,500 C. But the solar corona reaches millions of degrees. How is it possible to get so much hotter hundreds of thousands kilometers away from the Suns surface?

 Noted solar astronomer Eugene Parker for whom the Parker Solar probe was named suggested it was due to nanoflares small flares emanating from the solar surface much smaller than the usual solar flares:

ScienceCasts: The Mystery of Nanoflares.

 But what causes the nanoflares? Could it be asteroidal impacts? The argument could be made they are too small to cause any visible reaction on the Sun. But the question is of the local impact. The Sun’s escape velocity at its surface is 600 km/s. That is a tremendous amount of energy for a body impacting it at that speed. When material is thrown up after the impact the high temperature could be maintained far above the surface.

Nanoflares and coronal heating.

 Micro-flare observed on 4 September 2016 with NASA SDO/AIA and the Swedish 1-m Solar Telescope.

The image shows a micro-flare observed on 4 September 2016. Magnetic reconnection in the corona as sketched in the cartoon in the lower left produces a hot loop of more than 7 million degrees. This hot loop is visible as the bright area in the green background image taken with the Solar Dynamics Observatory (AIA 94 Å). The active region with bright magnetic loops is shown in more detail in the yellow inset, corresponding to plasma of less than 1 million degrees (AIA 171 Å). The reconnection event in the corona produces fast electrons that hit the lower atmosphere with high energy. The impact region is very small and is shown at high resolution in the image taken with the Swedish 1-m Solar Telescope on La Palma. With the European Solar Telescope, we will be able to study the magnetic environment of the impact region in even finer detail.

https://est-east.eu/?option=com_content&view=article&id=920&Itemid=622&lang=en

 For instance Jupiter’s escape velocity is 60 km/s and we saw the tremendous resulting impact from comet Shoemaker-Levy when it impacted Jupiter.

Jupiter in infrared, Shoemaker-Levy 9 collision (left) and Io (right) by Max Planck Institute for Astronomy  

 But the major, key reason for suspecting it is this: there is a type of nuclear fusion called impact fusion. It arises when bodies are made to collide at hundreds of kilometers per second relative impact speed. 

Proceedings of the
Impact Fusion Workshop ~ National Security and Resources Study Center
LOS Alamos Scientific Laboratory Los Alamos, New Mexico
LOSALAMOS SCIENTI
LABORATORY
PostOfficeBox 1663 LosAlamos,New Mexico87545
July 10—12, 1979

https://sgp.fas.org/othergov/doe/lanl/lib-www/la-pubs/00203259.pdf

There are private fusion research concerns now investigating this to bring about controlled nuclear fusion. 

 Recent observations of nanoflares have observed million degree temperatures locally around the nanoflares origin point on the Sun’s surface, while the surrounding area is at the normal 5,500 C temperature.

 So why don’t we see the asteroids during imaging of the nanoflares? It could be their small size. The Sun is so bright it completely washes out the asteroids that may be only a few kilometers across.

Recent observations and theoretical modeling suggest the million degree temperatures seen in the vicinity of the nanoflare origin point on the Sun’s surface should be able to be communicated to the corona-sphere thousands of kilometers above the Suns surface:

This May Be the First Complete Observation of a Nanoflare.
Heating the corona.
So far, these bright loops appeared to be tiny flares – but did their heat actually reach the corona?Bahauddin looked to NASA’s Solar Dynamics Observatory, which carries telescopes tuned to see the extremely hot plasma only found in the corona. Bahauddin located the regions right above the brightenings shortly after they appeared. “And there it was, just a 20-second delay,” Bahauddin said. “We saw the brightening, and then we suddenly saw the corona got super-heated to multi-million degree temperatures,” Bahauddin said. “SDO gave us this important information: Yes, this is indeed increasing the temperature, transferring energy to the corona.” Bahauddin documented 10 instances of bright loops with similar effects on the corona. Still, he hesitates to call them nanoflares. “Nobody actually knows because nobody has seen it before,” Bahauddin said. “It’s an educated guess, let’s say.”From the perspective of the theory that says nanoflares heat the corona, the only thing left to do is to show that these brightenings occur often enough, all over the Sun, to account for the corona’s extreme heat. That’s still work in progress. But observing these tiny bursts as they heat solar atmosphere is a compelling start.
https://www.nasa.gov/solar-system/this-may-be-the-first-complete-observation-of-a-nanoflare/

Additionally I was startled see to what would be the kinetic energy of an asteroid impacting the Sun at the 600 km/s escape velocity. Asteroids have been estimated to have densities in the range of 2,000 kg/m3 to 5,000 kg/m3 . The iron-nickel asteroids would have the higher density. This is important because they could also maintain their cohesiveness as they impacted the Sun.

Searches of a population of asteroids inside the orbit of Mercury, called vulcanoids, have been unsuccessful. This is because you have to look at the bright solar disk to detect them. But such searches put a size limit of 6 km wide on them. So assume the asteroid has size, say, 5 km across, with density, say, 4,000 kg/m3 . At that density the mass would be 4,000 kg/m^3 * (6,000 m)^3 = 8.64 * 10^14 kg. Now suppose this impacted the Sun at 600 km/s. Then the kinetic energy of that pact would be:

(1/2) * 8.64 * 10^14 * (600,000 m)^2 = 1.55*10^26 Joules. That is a tremendous amount of energy! To put in perspective the energy the Sun puts out each second is 3.86 * 10^26 watts. So if the asteroid deposited that energy in, say, 1 second, it would be a significant percentage of the total energy the Sun puts out in a second!

 However, asteroids of kilometers size impacting the Sun must be quite rare, judging from this graph of asteroid impacts to Earth by size:

https://www2.tulane.edu/~sanelson/Natural_Disasters/impacts.htm

 One meter and below must be more common. If the meteor impacting the Sun was 1 meter wide, then the kinetic energy would be (1/2)*4,000*(1)^3*(600000)^2 = 7.2*10^14 joules, nearly a quadrillion joules of energy.

 About the likelihood of asteroids impacting the Sun in accordance to the change needed in their established orbital velocity, if they started further out in the Solar System, much less velocity change (delta-v) would be needed to direct them to impact the Sun.

 Key confirmation required is to confirm the existence of these small solar impactors. Observations in the visual light spectrum have not succeeded. This is the solar irradiance spectrum showing the range of intensity’s according to wavelength:

  You see it is vanishingly small at extreme ultraviolet wavelengths and at radio wavelengths around 10,000 nm, 10 microns, and above. The problem with observations at the extreme ultraviolet is that not large enough telescopes have been launched to observe them at less than 6 km diameters (the extreme UV is absorbed by the Earth’s atmosphere.)

 Then the suggestion is to use large radio telescopes at the micron and above wavelengths to detect the close in asteroids.

 One radio telescope that might manage it is the ALMA radio telescope array:

ALMA Demonstrates Highest Resolution Yet

November 15, 2023 | Science

The Band-to-band (B2B) method demonstrated this time to achieve the highest resolution with ALMA. In the B2B method, atmospheric fluctuations are compensated for by observing a nearby calibrator in low frequency radio waves, while the target is observed with high frequency radio waves. The top right inset image shows the ALMA image of R Leporis that achieved the highest resolution of 5 milli-arcsec. Submillimeter-wave emissions from the stellar surface are shown in orange and hydrogen cyanide maser emissions at 891 GHz are shown in blue. The top left inset image shows a previous observation of the same star using a different array configuration with less distance between the antennas and without the B2B method, resulting in a resolution of 75 milli-arcsec. The previous resolution is too coarse to specify the positions of each of the two emission components. (Credit: ALMA (ESO/NAOJ/NRAO), Y. Asaki et al.) Download image (1.3MB)

https://www.nao.ac.jp/en/news/science/2023/20231115-alma.html

 At a max resolution of 5 milli-arc it should be able to detect kilometer wide asteroids at the distance of the Sun. The detection sensitivity should also be improved for iron-nickel meteorites for radio astronomy.

 Note the importance of this is that if it is confirmed then we know impact fusion does indeed work.

  Bob Clark