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Advanced Concepts / Re: Using shaped charges as a rocket engine
« Last post by Beratnyi on Today at 08:39 pm »
TNT has an energy density of 4.184 MJ / kg. Explosives apparently exist with up to 2.38 times greater energy density than TNT (https://en.wikipedia.org/wiki/TNT_equivalent), but those explosives have not been synthesized in any quantity, probably for good reasons, so I'll use 1.90, the highest value for an explosive that seems like someone actually seriously considered using. With perfect efficiency the best possible exhaust velocity is sqrt(2 * 1.90 * 4.184 MJ / kg) = 3,987 m/s. That's not bad, but that assumes 100% efficiency, i.e. that the products all leave the rocket in the desired direction at the same speed. I don't know how efficient shaped charges can be but my hunch is you'd get less than half of this, probably much less, which makes this idea not competitive with traditional rocket propellants. Two things are hurting this idea: the energy density of explosives isn't as good as for bipropellants, and the efficiency of shaped charges is probably much worse than a nozzle's efficiency.

Something similar has been suggested using nuclear bombs: https://en.wikipedia.org/wiki/Project_Orion_(nuclear_propulsion) .
Shaped charges typically use explosives with an energy density of 9 MJ/kg, which is comparable to methalox. But the most important thing is the gigantic pressure of hundreds of thousands of bars during the formation of a cumulative jet, which allows it to be accelerated to a speed of 10 km/s or more. For comparison, in the combustion chamber of the Raptor engine, the pressure is "only" 300 bar, and the jet velocity is 3.5 km / s, and this is an incredible achievement - it is not possible to significantly increase this value in a classic rocket engine.
A nozzle in a rocket is needed for only one task - the formation of a directed high-speed jet. In a shaped charge, this jet is formed by the charge itself, so no nozzle is needed.
To form a jet, a shaped charge is accelerating a tamper, not the charge itself*. That tamper is not part of the combustion process, so is pure dead reaction mass that reduces your effective MJ/kg (e.g. if you have 1kg of explosive accelerating a 1kg tamper into a EFP, then your energy density if halved).

Then you have the problem of harvesting that energy for propulsion. A single shaped charge floating in space may be able to fling a het of hot copper to a few km/s. but there's no vehicle that accelerates in the opposite direction in that scenario. You'd need to contain the explosion somehow and capture the gasses that are produces as a result of the explosion in order to use them to push a vehicle forwards. The problem there is that whilst a shaped charge is good at accelerating a tamper in one direction, it's no so good at accelerating the combustion products in one direction: instead, they expand mostly radially: you do not get the equivalent of a lump of gas travelling at the opposite speed to the tamper jet.
We don't need a metal liner! We are not going to penetrate the armor of the tank, all we need is a high-speed jet directed in the direction we need.
The liner forms the jet. No liner, no jet.
The first shaped charges did not have a metal liner, but they perfectly formed a directed jet of hot gases.
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Historical Spaceflight / Re: First Lunar Outpost (1992)
« Last post by Blackstar on Today at 08:28 pm »
This is a collection of material from one of the FLO workshops held in 1992. I'm not sure if this is everything I have. I'll have to dig around a bit more.
Nice reading material. Here is some more FLO material mostly via NTRS except the first.


Thanks for this collection. I cannot remember where I got my FLO material. I might have obtained it in Houston, or at NASA HQ.

See my TSR articles for some of the context for FLO. If you were to look at a history of NASA post-Apollo studies about sending humans to the Moon, you'd find very little before FLO. The big 90-Day Study done in 1989 was all done at JSC and appears to have been a "throw everything at the wall and see what sticks" effort. There were some truly crazy assumptions in there, like they were going to expand the VAB and (I think) build a third shuttle launch pad. It was not really focused. Their starting assumption was to try to do everything at once--start a Moonbase and a human mission to Mars. So the infrastructure costs were immense.

It took several years to get to a pretty basic starting point: what is the first step in returning humans to the Moon?

But even then, that starting point question had to include a number of assumptions, like how many astronauts, how long on the surface, what capabilities, etc.? Those starting assumptions were essentially to double Apollo: four astronauts, up to a month on the surface, the entire surface and not just the equator. That drove the FLO architecture towards a new Saturn V+ class rocket. And as a result, FLO was going to be very expensive and not use existing shuttle hardware.

After FLO, the lunar return plans were dead (they were dead before FLO too, but that's besides the point). Only then did some people start to ask if they could do a lunar return that accomplished less than FLO. What would it take to essentially re-do Apollo in terms of two astronauts on the Moon for a couple of days? Unfortunately, the options there were not good either.

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Have there been any studies done as of late on using the Vulcan rocket as an alternative to the SLS Core Booster to launch the Orion capsule in a distributed launch profile?

Boeing (which co-owns ULA with Lockheed Martin) periodically orders ULA not to do or say anything that could threaten SLS (which Boeing has a big contract to develop), e.g. https://arstechnica.com/science/2019/08/rocket-scientist-says-that-boeing-squelched-work-on-propellant-depots/. That's probably why ULA isn't proposing SLS replacements any more.
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But why only 40? Given the dimensions, 56 can probably also be placed. The Soyuz has much less space under the fairing, and the Fregat reduces it even more, but it placed a block of the same height, albeit a truncated one. Perhaps this is what Oneweb was able to give.
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Quote
Huge!

https://twitter.com/nextspaceflight/status/1599827071060283392

Quote
Blue Origin is wrapping up taking a New Glenn payload fairing halve for a swim at the KSC turn basin this morning. It appears that Blue plans to recover the fairing halves from the water similar to SpaceX.

nsf.live/spacecoast

But the New Glenn fairing is missing one part of the Falcon fairing "magic sauce" for ocean surface recovery.
The Falcon fairing has a larger diameter than the second stage.
That means that the back end, where it attaches to the second stage, necks down.
This gives the fairing halves the shape of a boat.
The New Glenn fairing is the same diameter as the rocket.
That leaves the back end open.
It is shown this way in Blue's drawings of the fairing in flight, where there are graphic cut-outs to see the payload.
That makes this fairing unsuitable for floating, unless they add something.
Perhaps an inflatable transom?
Maybe just a truncated half cone around the payload adapter.
Notice that the posted photos do not show the back end of the fairing being lifted by the crane.

It is still incomprehensible that anyone would lable this "Huge!"
Blue has a long, long way to go, and even "following" by copying SpaceX will have daunting challenges.
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New paper published today from InSight data shows that Mars may still be more active than previously suspected.

Geophysical evidence for an active mantle plume underneath Elysium Planitia on Mars

Abstract
Although the majority of volcanic and tectonic activity on Mars occurred during the first 1.5 billion years of its geologic history, recent volcanism, tectonism and active seismicity in Elysium Planitia reveal ongoing activity. However, this recent pulse in volcanism and tectonics is unexpected on a cooling Mars. Here we present observational evidence and geophysical models demonstrating that Elysium Planitia is underlain by an ~4,000-km-diameter active mantle plume head. Plume activity provides an explanation for the regional gravity and topography highs, recent volcanism, transition from compressional to extensional tectonics and ongoing seismicity. The inferred plume head characteristics are comparable to terrestrial plumes that are linked to the formation of large igneous provinces. Our results demonstrate that the interior of Mars is geodynamically active today, and imply that volcanism has been driven by mantle plumes from the formation of the Hesperian volcanic provinces and Tharsis in the past to Elysium Planitia today.

https://www.nature.com/articles/s41550-022-01836-3
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Just Kerbal it and stick some SRBs in the skirt for pad aborts.

You joke, but there is room for 3 solid rocket boosters on Starship with a burn time of a few seconds.   Instead of adding 3 vacuum raptors, add 3 solid rocket boosters of the same dimensions.

The joke is probably applicable.  First of all, this is a new engine.  SRBs aren't particularly complicated, but still, if you're comparing the amount of work to something like the enshrouded D2 escape system, they're within shouting distance of one another.

Also, this makes the mass-flow problem in the skirt worse.  If you're going to start any engines on the Starship while still mated to the SH, you're going to have to engineer blowout panels at the very least.  But at some point, there's so much mass flow, irrespective of blowout panels, that you're going to destroy the top of the SH LCH4 dome, which will be game over.  Since SRBs derive their thrust from insane amounts of mass flow, they'll make things a lot worse.

A quick note:  The T/W=6 with the most lightly loaded crewed Starship possible to orbit already assumes 9 engines.

If solid rockets don't work then there is no pad abort for a Starship loaded up to 1400t with fuel.  your same arguments apply to having a crew dragon in the cargo compartment.

So if there is no pad abort, and ascent abort works with current design, and there is no ELD abort, I'm flummoxed as to what abort system there needs to be at all.

The above 3 stages of flight are the highest risk stages.

Landing abort is all that's left.   "can't get to catch tower" means water abort, which works with a robust cargo compartment.

Failure to ignite or correctly use engines due to the remaining single points of failure (which is tank pressurization and gimbaling) happens at such a low altitude abort would be difficult.   Blasting a crew dragon horizontally won't be useful, and neither will solid rockets in the base.

I suspect gimbaling can be made redundant barring frozen parts1, so really we are just left with inadequate pressure in the header tanks as the remaining single point of failure.



(1) There are still single points of failures in commercial airplanes, e.g. the jack screw on the 737 which has killed people
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The most basic item after [re-usability] is being able to do propellant transfer.
With those 2 pieces in place SX will send something Starship sized to Mars in 2024.

Maybe the will maybe the won't.

We certainly can't be certain! But unless something changes in the underlying motivations of the enterprise, we can be fairly confident they will want to keep Mars on their agenda.

[Re-usability] is nice to have. Even, it's a requirement for a long term solution. But it's not part of MVP (minimum viable product)
[and] propellant transfer is not required, either.

There's possibly a business case for SS/SH as a low cost, partially-expended, non-refilled launch system. SpaceX hasn't seemed interested in addressing that market a second time. Falcon is doing that just fine....
This wouldn't be about a business case, it would be about a one-off expendable mission to get some PR plus maybe a bit of deep-space experience with Starship, if that's all they can do by then.
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The most basic item after [re-usability] is being able to do propellant transfer.
With those 2 pieces in place SX will send something Starship sized to Mars in 2024.

Maybe the will maybe the won't.

We certainly can't be certain! But unless something changes in the underlying motivations of the enterprise, we can be fairly confident they will want to keep Mars on their agenda.

[Re-usability] is nice to have. Even, it's a requirement for a long term solution. But it's not part of MVP (minimum viable product)
[and] propellant transfer is not required, either.

There's possibly a business case for SS/SH as a low cost, partially-expended, non-refilled launch system. SpaceX hasn't seemed interested in addressing that market a second time. Falcon is doing that just fine....
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