Author Topic: Testing upper stage propulsive entry survivability  (Read 9523 times)

Online Kaputnik

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Thinking about how SpaceX approached the issue of first stage recovery testing, is there any merit in doing the same with the second stage? That is, use margin on an operational launch, using existing hardware, to see how far through entry the vehicle can get.

Using the numbers of Ed Kyle's site, I get a delta V of about 11,000m/s for the second stage.
If there were no payload at all (hypothetical scenario), I estimate you would have around 5,300m/s to deorbit and slow the entry. That's quite a lot, but is it enough?

We know what combination of speed/altitude the first stage can survive, but I lack the knowledge to translate that information into the relevant conditions for the upper stage.

Secondly, does it seem likely that the second stage might be more robust than the first, from an entry standpoint?

So my main question is... does the forum think that there is anything to be learned from attempting partial propulsive entry testing? And are there any particularly small payloads manifested that would be potential opportunities (like CASIOPE was).
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Offline Lar

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Re: Testing upper stage propulsive entry survivability
« Reply #1 on: 02/20/2017 04:28 PM »
I think the theory is the numbers miss by so much that there's no merit in trying to actually recover. But you may be on to something in that there may be things that can be learned by higher speed reentries even if the stage isn't recovered. But what? What does SpaceX need to know that it hasn't gotten from S1 reentry and the low density heat shield work NASA already did?

If we can spot something?
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Offline envy887

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Re: Testing upper stage propulsive entry survivability
« Reply #2 on: 02/20/2017 05:30 PM »
It would be interesting to see if (or how long) the vac nozzle can survive hypersonic entry aero loads. Even after launching a very light LEO payload and propulsive deceleration it will still be travelling somewhere around Mach 10 at entry interface.

Online Kaputnik

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Re: Testing upper stage propulsive entry survivability
« Reply #3 on: 02/20/2017 05:46 PM »
Where I would see the value in this experiment is in determining the survival point of the stage. Start the entry burn just before you think the stage is about to be lost, and make a long, throttled back burn to see how long the stage can survive. There is still a lot to be learned in hypersonic retropropulsion.
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Online whitelancer64

Re: Testing upper stage propulsive entry survivability
« Reply #4 on: 02/20/2017 05:54 PM »
Recovering the second stage with retropropulsion isn't going to work, too much fuel is needed. Just stick a heat shield (perhaps expandable) on the nose.
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Offline R7

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Re: Testing upper stage propulsive entry survivability
« Reply #5 on: 02/20/2017 07:17 PM »
S2 with proper heat shield doesn't even have to do atmospheric entry burn S1 does.
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Offline S.Paulissen

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Re: Testing upper stage propulsive entry survivability
« Reply #6 on: 02/22/2017 04:10 AM »
S2 with proper heat shield doesn't even have to do atmospheric entry burn S1 does.

Is there evidence for this assertion?  How much would the shield mass?  Is that factored into sim?  What are peak breaking gees?  Just curious.
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Offline R7

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Re: Testing upper stage propulsive entry survivability
« Reply #7 on: 02/22/2017 08:31 AM »
1. Is there evidence for this assertion? Is that factored into sim?
2. How much would the shield mass?
3. What are peak breaking gees?

1. Rational reasoning. Reusable S2 has to cope with heat and stresses while decelerating from orbital speed. Heat shield is given, more on this below, so whatever thermal environment the S1 couldn't cope without powered breaking was, S2 can do it. S2 is more stubby which adds to structural strength. And it has to be aerodynamically stable during reentry. Didn't early F9 S1 reentry attempts fail mainly because there were no means to control the attitude, no RCS no fins, so it didn't stabilize and kept tumbling until aerodynamic forces broke it?

2. There are calculations in other threads estimating Dragon heatshield weighing little over 200kg. Area is known and PICA-X density is known, it's very light stuff. Then there's of course some additional support structure but it's still a total of how many hundreds of kgs, not tons. Relying on that mass to decelerate from orbital speed to terminal velocity vs. hypersonic retropropulsion isn't much of a contest.

3. Hard to say for sure because need to know things like S2 ballistic coefficient, drag coefficient and whether it uses any lift or does purely ballistic reentry. BC would likely to be smaller than any returning SCs so far. IIRC low BC means greater deceleration in upper atmosphere but lower peak deceleration because velocity when entering the quickly thickening lower atmosphere is smaller. Read about this in some paper analysing ICBM RV trajectories, fun!  ;)

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Offline dorkmo

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Re: Testing upper stage propulsive entry survivability
« Reply #8 on: 02/23/2017 01:10 AM »
would the center of gravity be a problem?

Offline envy887

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Re: Testing upper stage propulsive entry survivability
« Reply #9 on: 02/23/2017 03:27 AM »
S2 will have a ballistic coefficient similar to Apollo unless it does a lifting entry, in which case the heat shield needs to extend down one side but the BC can be 4 or 5 times less.

With chines, a full length shield, and an engine flap it could enter like ITS, with a very high alpha, lots of lift, and quite low ballistic coefficient.

Offline corneliussulla

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Re: Testing upper stage propulsive entry survivability
« Reply #10 on: 02/23/2017 10:00 AM »
The ITS is effectively a 2nd stage re entry vehicle. A similar type vehicle could be built for launch on falcon I would think. In a world where ITS refuelling exists refuelling a notional falcon 2nd stage for deorbit would be a minor affair I would think. ITS tankers could be parked in orbit to enable a great number of falcon 2nd stages to return to earth for each ITS tanker launch. If ITS tanker will be available in next 6 years this would seem the most logical way forward rather than trying to make the existing 2nd stage survivable

Re: Testing upper stage propulsive entry survivability
« Reply #11 on: 02/23/2017 09:52 PM »
Seems like a good thread to re-post this old 2011 animation showing S2 re-entry, landing. But would it be able to land back at the cape? Would it have to orbit first?


Offline speedevil

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Re: Testing upper stage propulsive entry survivability
« Reply #12 on: 02/24/2017 02:15 AM »
Seems like a good thread to re-post this old 2011 animation showing S2 re-entry, landing. But would it be able to land back at the cape? Would it have to orbit first?
Second engine cutoff is around 600 seconds into flight, at ~8km/s. Neglecting stuff, that's probably over 1/3*8*600 = 1500km or so downrange.
In order to get back to the takeoff site without orbiting, you're going to need to do something wacky - as you are at this point in orbit. Something like a quite substantial burn to intersect the atmosphere, use the atmosphere to kill most of the velocity, burn back up ballistically to the takeoff site, entry trim burns perhaps, and then entry and landing burns like F9 S1.

This is going to need a _lot_ more delta-v than waiting around a day, perhaps with a trim manoever to adjust phasing and going straight in for a landing without the intervening~1500km ballistic hop needing several kilometers/s worth of fuel.
Unless of course you add high performance wings.
(which would be insane)

Offline Lars-J

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Re: Testing upper stage propulsive entry survivability
« Reply #13 on: 02/24/2017 05:41 PM »
Seems like a good thread to re-post this old 2011 animation showing S2 re-entry, landing. But would it be able to land back at the cape? Would it have to orbit first?

Short answer: A reusable 2nd stage could only make it back to the launch site after one orbit. (it might have to wait more orbits, perhaps as long as a day)

Offline mvpel

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Re: Testing upper stage propulsive entry survivability
« Reply #14 on: 02/24/2017 09:37 PM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
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Offline Jcc

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Re: Testing upper stage propulsive entry survivability
« Reply #15 on: 02/26/2017 01:24 PM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.

They have clearly determined that there is no way to make it work for Falcon, but it is already an integral feature for ITS. It works (hopefully) with ITS by scaling to 10x more thrust in the first stage and over 20x more thrust in the 2nd stage/payload, on orbit refueling, all composite structure, low bulk density for the nearly empty returning vehicle, etc.  you could call these non-trivial changes.

Offline deruch

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Re: Testing upper stage propulsive entry survivability
« Reply #16 on: 03/08/2017 09:35 PM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together. 

edit: Both mattstep and saliva_sweet have subsequently pointed out a serious flaw in my explained reasoning (seven posts further down).  Quite an embarrassing mistake.  As a result, of this and in conjunction with Elon's continued, recent statements about the comparative costs of each stage, I have since revised my estimate downwards to 1MVac = 3-4 M9.  Sorry for the delayed edit.
« Last Edit: 04/16/2017 03:23 AM by deruch »
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Offline S.Paulissen

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Re: Testing upper stage propulsive entry survivability
« Reply #17 on: 03/09/2017 04:06 AM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together.

Perhaps your reputation alone can attest to this, but I would like to see this claim substantiated with some evidence. 
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Offline Danderman

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Re: Testing upper stage propulsive entry survivability
« Reply #18 on: 03/09/2017 04:24 AM »
I vaguely remember a Gary Hudson concept where the upper stage would return engine first, with the engine firing during re-entry. Apparently, this would be sufficient for the stage to survive the re-entry heating.

Online JamesH65

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Re: Testing upper stage propulsive entry survivability
« Reply #19 on: 03/09/2017 10:10 AM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together.

Perhaps your reputation alone can attest to this, but I would like to see this claim substantiated with some evidence.

Indeed - I've not seen any claims or data on this before.

Offline mvpel

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Re: Testing upper stage propulsive entry survivability
« Reply #20 on: 03/09/2017 12:18 PM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together.

Per the public statements of Elon Musk:

Quote from: Elon Musk via Motley Fool
As Elon Musk explains, "The boost stage [of a Falcon 9 rocket] is about 70% of the cost of the rocket ... it's sort of on the order of $30 to $35 million dollars."

It seems far-fetched to me that the Mvac would cost in excess of 9x more than the SL Merlin when those 9 Merlins plus legs, fins, and some longer tanks rolled out with the same equipment as the upper stage tanks is 70% of the $43 - $50 million cost of the rocket.
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Online Kaputnik

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Re: Testing upper stage propulsive entry survivability
« Reply #21 on: 03/09/2017 08:51 PM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together. 

Niobium must be really expensive stuff.

Source?
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Online HMXHMX

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Re: Testing upper stage propulsive entry survivability
« Reply #22 on: 03/09/2017 10:42 PM »
I vaguely remember a Gary Hudson concept where the upper stage would return engine first, with the engine firing during re-entry. Apparently, this would be sufficient for the stage to survive the re-entry heating.


My approach was a bit more subtle than stated here; my idea dates from the early 1980s when I was looking at recovery of a lifting-ballistic VTOL, base-first, and proposed firing a small centrally-located thruster or gas generator to push off the main re-entry shock.  That dramatically lowers the conductive, radiative and convective heating of the base.  Applied to the F9 S2, it might be possible to operate the engine's GG only, perhaps with additional H2O injected to increase mass flow and to lower temperatures, and then to exhaust the flow out of the main bell.  But the large bell would almost certainly need to be jettisoned first.  This concept is informed by some of the wind tunnel work done by NASA Langley on supersonic retropropulsion, which can be seen on youtube videos.

Offline deruch

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Re: Testing upper stage propulsive entry survivability
« Reply #23 on: 03/11/2017 12:27 AM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together.

Perhaps your reputation alone can attest to this, but I would like to see this claim substantiated with some evidence.
Was based on a series of comments from a former SpaceX employee who worked on the MVac.  @gongora put them all together in a comment in the Merlin 1D thread (quoted pertinent bits below).  Most specifically that at the time he worked there, it took 1-2 days to produce an M1D vs. 18-21 days for an MVac. due to its complexity.  My understanding of the disparate costs wasn't based on materials cost but man-hours.

The Reddit comments section is here: https://www.reddit.com/r/spacex/comments/5h94xv/picture_of_a_mvac_engine_sitting_inside_its/

Some interesting notes from a former employee who used to work on the MVac engines.  He left in the Fall of 2015 so some things may be a little out of date.
Comment by Foximus05 on whether employees floated between different tasks:
Quote
Not when I was there. You might float if someone was behind and needed help, but there was a set tam that only did MVAC, only did M1D's, only did octaweb, etc. the M1D guys might move around from lowers to uppers, or chambers, but not Mvac, because it required so much more attention to detail.
Comment by Foximus05 on assembly times for the engines:
Quote
When I left it was a day or two for an M1D (dependant on parts) Vs 18-21 days for an MVAC. Mvac is a lot more complex, has more systems and has a bunch of made on assembly parts
Comment by Foximus05 on M1D vs. MVac
Quote
Very. MVAC contains more systems that M1D's have inside the octaweb, along with some control valves for the second stage. The chamber and a few other parts are the only similarities. Its in the same class, but its like comparing a Small Block Chevy V8 to a Ferrari engine.
Comment by Foximus05 on M1D vs. MVac:
Quote
MVAC was around ~400 pounds heavier than M1D, sans vacuum nozzle. But that was before they went to FT and I left.
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Offline stcks

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Re: Testing upper stage propulsive entry survivability
« Reply #24 on: 03/11/2017 12:54 AM »
Comment by Foximus05 on M1D vs. MVac
Quote
Very. MVAC contains more systems that M1D's have inside the octaweb, along with some control valves for the second stage. The chamber and a few other parts are the only similarities. Its in the same class, but its like comparing a Small Block Chevy V8 to a Ferrari engine.

Pardon the stupid question, but is this common for upper stage engines in general? Or is this just unique to MVac vs M1D? What are the reasons why MVac should be so much more complicated than M1D?

Offline mattstep

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Re: Testing upper stage propulsive entry survivability
« Reply #25 on: 03/11/2017 01:23 AM »
I can certainly see why they've tabled this for the time being - they're already recovering 90% of the engines and 70% of the cost of the launch vehicle, and bringing down that last engine and its tankage is clearly a very tricky proposition.
1 MVac costs more than all 9 of the SL Merlins put together.

Perhaps your reputation alone can attest to this, but I would like to see this claim substantiated with some evidence.
Was based on a series of comments from a former SpaceX employee who worked on the MVac.  @gongora put them all together in a comment in the Merlin 1D thread (quoted pertinent bits below).  Most specifically that at the time he worked there, it took 1-2 days to produce an M1D vs. 18-21 days for an MVac. due to its complexity.  My understanding of the disparate costs wasn't based on materials cost but man-hours.

The Reddit comments section is here: https://www.reddit.com/r/spacex/comments/5h94xv/picture_of_a_mvac_engine_sitting_inside_its/

Some interesting notes from a former employee who used to work on the MVac engines.  He left in the Fall of 2015 so some things may be a little out of date.
Comment by Foximus05 on whether employees floated between different tasks:
Quote
Not when I was there. You might float if someone was behind and needed help, but there was a set tam that only did MVAC, only did M1D's, only did octaweb, etc. the M1D guys might move around from lowers to uppers, or chambers, but not Mvac, because it required so much more attention to detail.
Comment by Foximus05 on assembly times for the engines:
Quote
When I left it was a day or two for an M1D (dependant on parts) Vs 18-21 days for an MVAC. Mvac is a lot more complex, has more systems and has a bunch of made on assembly parts
Comment by Foximus05 on M1D vs. MVac
Quote
Very. MVAC contains more systems that M1D's have inside the octaweb, along with some control valves for the second stage. The chamber and a few other parts are the only similarities. Its in the same class, but its like comparing a Small Block Chevy V8 to a Ferrari engine.
Comment by Foximus05 on M1D vs. MVac:
Quote
MVAC was around ~400 pounds heavier than M1D, sans vacuum nozzle. But that was before they went to FT and I left.

I don't think we have enough information here to make a cost comparison. There is nothing in those quotes about numbers of people on each team to make a comparison on total labor hours to produce each engine.

Offline Adaptation

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Re: Testing upper stage propulsive entry survivability
« Reply #26 on: 03/11/2017 08:24 AM »
"Really tempting to redesign upper stage for return too (Falcon Heavy has enough power), but prob best to stay focused on the Mars rocket" -Elon

https://twitter.com/elonmusk/status/755167487017291776

SpaceX is also experimenting with an upper stage raptor for F9/FH.

https://www.defense.gov/News/Contracts/Contract-View/Article/642983

There is no real rush on this, core block 5 will be more or less finalized but new upper stages will be needed continually for the foreseeable future.  A reusable version could be developed years down the line.   It would likely be raptor based and still be expendable for F9 the reusable version would probably only fly on FH.

The current Merlin sized raptor was mostly 3d printed so perhaps machining and manufacture time wont be such a big deal.  Or maybe if the Vulcan reuse system of just returning the engines with an inflatable heat shield works well SpaceX could copy that for the upper stage.

Offline saliva_sweet

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Re: Testing upper stage propulsive entry survivability
« Reply #27 on: 03/11/2017 09:55 AM »
it took 1-2 days to produce an M1D vs. 18-21 days for an MVac. due to its complexity.  My understanding of the disparate costs wasn't based on materials cost but man-hours.

That's flawed because you assume the same number of people working on both. It would make sense for them to proprtion their workforce such that they can produce 9 M1Ds during the time it takes to make an MVac.

Online Kaputnik

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Re: Testing upper stage propulsive entry survivability
« Reply #28 on: 03/11/2017 11:56 AM »
Of course an Mvac *must* function flawlessly every time, whereas there is some margin for failure on the other engines. Would enhanced QC play a part?

Instinctively, though, it seems likely that Mvac is simply produced in smaller numbers by a smaller team.
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Offline uhuznaa

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Re: Testing upper stage propulsive entry survivability
« Reply #29 on: 03/31/2017 02:26 PM »
Recovering the second stage with retropropulsion isn't going to work, too much fuel is needed. Just stick a heat shield (perhaps expandable) on the nose.

A heat shield on the nose won't help at all. The CoG is very far back (engine, thrust structure), with nothing but empty tankage on the top. So the stage WILL fly with the nozzle forward as soon as it hits the atmosphere. Also the big vacuum nozzle won't keep together with that kind of aerodynamic loads. The only way to do this would be to have a heat shield at the bottom that covers the base of the nozzle, then jettison most of the nozzle (although everything needed to do this would interfere with radiation cooling of the nozzle during ascend), do a long throttled reentry burn with lots of fuel kept for this and hope for the best.

And even then the engine would be much too powerful to land on it. Well, maybe with a really brutal suicide burn. But probably you'd need dedicated landing engines. And legs of course. And steerable fins for controlling the descent. And enough power for the time you have to wait until you again arrive over the landing site in orbit. And now you will have ended up with almost no payload.

The F9 second stage is just too small to be made reusable and still serve roughly the same payload market. You need a bigger craft to do that (even if not as big as ITS).


Online guckyfan

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Re: Testing upper stage propulsive entry survivability
« Reply #30 on: 03/31/2017 02:54 PM »
A heat shield on the nose won't help at all. The CoG is very far back (engine, thrust structure), with nothing but empty tankage on the top. So the stage WILL fly with the nozzle forward as soon as it hits the atmosphere. Also the big vacuum nozzle won't keep together with that kind of aerodynamic loads. The only way to do this would be to have a heat shield at the bottom that covers the base of the nozzle, then jettison most of the nozzle (although everything needed to do this would interfere with radiation cooling of the nozzle during ascend), do a long throttled reentry burn with lots of fuel kept for this and hope for the best.

And even then the engine would be much too powerful to land on it. Well, maybe with a really brutal suicide burn. But probably you'd need dedicated landing engines. And legs of course. And steerable fins for controlling the descent. And enough power for the time you have to wait until you again arrive over the landing site in orbit. And now you will have ended up with almost no payload.

The F9 second stage is just too small to be made reusable and still serve roughly the same payload market. You need a bigger craft to do that (even if not as big as ITS).

You believe Elon Musk has not thougt of these things?

Online Kaputnik

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Re: Testing upper stage propulsive entry survivability
« Reply #31 on: 03/31/2017 02:59 PM »
My premise in starting this thread was that on low performance missions, SpaceX will end up with a second stage with considerable residual propellant. So much in the same way as their initial tests with first stages, this paid-for opportunity could be exploited to research how well a rocket stage can survive entry when using a long, low engine burn to cushion the entry. There are quite large holes in our knowledge of how this would work, and it is of great interest to NASA as well as SpaceX, as it could be directly applicable to Mars entry. I wasn't really suggesting that this would allow recovery of the existing design of second stage, but I thought for the relatively low cost it would be valuable.

if we were to consider major hardware changes.... well then a front mounted heatshield and a toroidal aft inflatable skirt to change the CoP would be one way of doing things... perhaps that skirt would itself act as enough of a drogue to allow mid air recovery of the falling stage lower in the atmosphere? But there are other threads for that.
Waiting for joy and raptor

Online TrevorMonty

Re: Testing upper stage propulsive entry survivability
« Reply #32 on: 03/31/2017 07:29 PM »
The shape and flight profile of reusable 2nd stage likely be same as for ITS.

Offline Lar

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Re: Testing upper stage propulsive entry survivability
« Reply #33 on: 03/31/2017 07:38 PM »
Good bump on this older thread, based on Elon's "hail Mary" comment, perhaps they are going to try some of the experiments Kaputnik outlined on a mission that has low mass so there is spare propellant.  Would this also require adding some lifetime to the S2 to allow multiple orbits before starting down? I guess if you're just experimenting, you start burning whererever convenient, as long as your entry point is the South Pacific somewhere?

See also  http://forum.nasaspaceflight.com/index.php?topic=42637 which is similar but different focus.
« Last Edit: 03/31/2017 07:56 PM by Lar »
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
"We're a little bit like the dog who caught the bus" - Musk after CRS-8 S1 successfully landed on ASDS OCISLY

Online TrevorMonty

Re: Testing upper stage propulsive entry survivability
« Reply #34 on: 03/31/2017 07:44 PM »
Elon Musk (@elonmusk) tweeted at 7:44 AM on Sat, Apr 01, 2017:
Considering trying to bring upper stage back on Falcon Heavy demo flight for full reusability. Odds of success low, but maybe worth a shot.

Elon comment 2nd reuse wasn't after thought, looks like they have been working on it.

Economics of F9, FH  reusable 2nd stage maybe marginal but its lot cheaper than using ITS to iron out bugs. For ITS it has to work 1st time, every time.


« Last Edit: 03/31/2017 07:50 PM by TrevorMonty »

Offline uhuznaa

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Re: Testing upper stage propulsive entry survivability
« Reply #35 on: 04/02/2017 05:05 PM »
A heat shield on the nose won't help at all. The CoG is very far back (engine, thrust structure), with nothing but empty tankage on the top. So the stage WILL fly with the nozzle forward as soon as it hits the atmosphere. Also the big vacuum nozzle won't keep together with that kind of aerodynamic loads. The only way to do this would be to have a heat shield at the bottom that covers the base of the nozzle, then jettison most of the nozzle (although everything needed to do this would interfere with radiation cooling of the nozzle during ascend), do a long throttled reentry burn with lots of fuel kept for this and hope for the best.

And even then the engine would be much too powerful to land on it. Well, maybe with a really brutal suicide burn. But probably you'd need dedicated landing engines. And legs of course. And steerable fins for controlling the descent. And enough power for the time you have to wait until you again arrive over the landing site in orbit. And now you will have ended up with almost no payload.

The F9 second stage is just too small to be made reusable and still serve roughly the same payload market. You need a bigger craft to do that (even if not as big as ITS).

You believe Elon Musk has not thougt of these things?

Of course he thought of these things, that's the reason he didn't try to return the F9 second stage. It probably would eat half of the payload.

Now, with the Falcon Heavy things look different, it has plenty of payload to spare. The easiest way to return the second stage would be to add some ballast on top of the second stage, with a heat shield, more batteries, propellant tanks, some SuperDracos and legs, and some flaps on the bottom of the stage. This would move the CoG of the S2 with empty tanks forward so it could do a head-first reentry, maybe with a reentry burn with deeply throttled SuperDracos to move the shock wave away. Later control the trajectory with the flaps, for landing deploy the legs and land on the SuperDracos. This would add some tons to the second stage and take this out of the payload, but the Heavy has more than enough of that.

For FH missions that don't need all of the payload you could return the second stage this way, this wouldn't actually be much different from Dragon 2 with powered landings, the dry mass and diameter of both is pretty similar. You'd basically add the bottom part of a Dragon 2 to the top of the second stage and return and land it on its head.

For an experiment on the FH demo flight you could just try to add some ballast and a heat shield to the top of the second stage and try to make it survive reentry.

Offline watermod

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Re: Testing upper stage propulsive entry survivability
« Reply #36 on: 04/02/2017 11:59 PM »
Another way of looking at it would be a much cheaper disposable S2.
Could a very cheap disposable S2 be made with a small 3D printed Raptor and alternate materials to aluminium?

Offline SweetWater

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Re: Testing upper stage propulsive entry survivability
« Reply #37 on: 04/03/2017 12:33 AM »
Another way of looking at it would be a much cheaper disposable S2.
Could a very cheap disposable S2 be made with a small 3D printed Raptor and alternate materials to aluminium?

I don't think a Raptor-powered upper stage for Falcon 9 is going to make anything cheaper - or smaller, for that matter. Vacuum-optimized Raptor has been estimated to have a ~4 meter nozzle (source: https://www.nasaspaceflight.com/2016/10/its-propulsion-evolution-raptor-engine/ ); as it is, Falcon 9 has a 3.7 meter diameter.

Moving to a Raptor-powered upper stage would mean developing a new, wider-diameter upper stage. That would require a new, wider fairing to fit on the wider upper stage, a new interstage between S1 and S2, and modifying the strongbacks for the wider interstage/S2/fairing combination. That's in addition to all the pad modifications needed to support the methane fuel for a Raptor S2.

I don't think a Raptor-powered upper stage is out of the question for Falcon 9 at some point in the future, but it won't be for a while, and it will be neither small nor cheap.

Offline Rocket Surgeon

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Re: Testing upper stage propulsive entry survivability
« Reply #38 on: 04/03/2017 01:13 AM »
"Really tempting to redesign upper stage for return too (Falcon Heavy has enough power), but prob best to stay focused on the Mars rocket" -Elon

https://twitter.com/elonmusk/status/755167487017291776

SpaceX is also experimenting with an upper stage raptor for F9/FH.

https://www.defense.gov/News/Contracts/Contract-View/Article/642983

There is no real rush on this, core block 5 will be more or less finalized but new upper stages will be needed continually for the foreseeable future.  A reusable version could be developed years down the line.   It would likely be raptor based and still be expendable for F9 the reusable version would probably only fly on FH.

The current Merlin sized raptor was mostly 3d printed so perhaps machining and manufacture time wont be such a big deal.  Or maybe if the Vulcan reuse system of just returning the engines with an inflatable heat shield works well SpaceX could copy that for the upper stage.

A few questions about this:
Has there been any word on this Raptor Upper Stage for the Falcon 9?
Wasn't congress trying to kill this or something?
Does the contract say that SpaceX actually has to build and fly it or is it just paper work and component testing?

Offline meekGee

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Re: Testing upper stage propulsive entry survivability
« Reply #39 on: 04/03/2017 01:27 AM »
Recovering the second stage with retropropulsion isn't going to work, too much fuel is needed. Just stick a heat shield (perhaps expandable) on the nose.

A heat shield on the nose won't help at all. The CoG is very far back (engine, thrust structure), with nothing but empty tankage on the top. So the stage WILL fly with the nozzle forward as soon as it hits the atmosphere. Also the big vacuum nozzle won't keep together with that kind of aerodynamic loads. The only way to do this would be to have a heat shield at the bottom that covers the base of the nozzle, then jettison most of the nozzle (although everything needed to do this would interfere with radiation cooling of the nozzle during ascend), do a long throttled reentry burn with lots of fuel kept for this and hope for the best.

And even then the engine would be much too powerful to land on it. Well, maybe with a really brutal suicide burn. But probably you'd need dedicated landing engines. And legs of course. And steerable fins for controlling the descent. And enough power for the time you have to wait until you again arrive over the landing site in orbit. And now you will have ended up with almost no payload.

The F9 second stage is just too small to be made reusable and still serve roughly the same payload market. You need a bigger craft to do that (even if not as big as ITS).
There's a leap in logic between where you finish counting the difficulties (and your solutions), and where you conclude that no reasonable payload capacity is left.

You have 20 tons of payload to start with.

How much do you figure are:
The heat shield
The parachute
The propellant reserve?

Assume air capture for a moment.
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Offline push2eject

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Re: Testing upper stage propulsive entry survivability
« Reply #40 on: 04/03/2017 01:33 AM »
Recovering the second stage with retropropulsion isn't going to work, too much fuel is needed. Just stick a heat shield (perhaps expandable) on the nose.

Hi all; my first post here. I don't know the math, but it seems like it shouldn't be too hard to work out exactly how much fuel is needed - for a purely propulsive de-orbit & decent, without a significant heatshield, to a point where a parachute can be deployed? Is this within the realm of possibility for the FH demo?

Offline meekGee

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Re: Testing upper stage propulsive entry survivability
« Reply #41 on: 04/03/2017 01:43 AM »
Recovering the second stage with retropropulsion isn't going to work, too much fuel is needed. Just stick a heat shield (perhaps expandable) on the nose.

A heat shield on the nose won't help at all. The CoG is very far back (engine, thrust structure), with nothing but empty tankage on the top. So the stage WILL fly with the nozzle forward as soon as it hits the atmosphere. Also the big vacuum nozzle won't keep together with that kind of aerodynamic loads. The only way to do this would be to have a heat shield at the bottom that covers the base of the nozzle, then jettison most of the nozzle (although everything needed to do this would interfere with radiation cooling of the nozzle during ascend), do a long throttled reentry burn with lots of fuel kept for this and hope for the best.

And even then the engine would be much too powerful to land on it. Well, maybe with a really brutal suicide burn. But probably you'd need dedicated landing engines. And legs of course. And steerable fins for controlling the descent. And enough power for the time you have to wait until you again arrive over the landing site in orbit. And now you will have ended up with almost no payload.

The F9 second stage is just too small to be made reusable and still serve roughly the same payload market. You need a bigger craft to do that (even if not as big as ITS).
There's a leap in logic between where you finish counting the difficulties (and your solutions), and where you conclude that no reasonable payload capacity is left.

You have 20 tons of payload to start with.

How much do you figure are:
The heat shield
The parachute
The propellant reserve?

Assume air capture for a moment.

Quoting myself..

The other thing to consider is that LEO missions are often volume limited, so the upper stage doesn't need its full payload capacity.

So let go the GEO missions.  Focus on LEO, and suddenly losing 30$ of the payload, for example, is not the end of the world.

Make your heat shield a "mission kit", and you basically have two flavors of the second stage - "reusable", and "full capability". 

ABCD - Always Be Counting Down

Offline Elmar Moelzer

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Re: Testing upper stage propulsive entry survivability
« Reply #42 on: 04/03/2017 04:36 AM »
I am not sure they will want to recover it after GEO missions. Those are already challenging enough for first stage reuse even. But for missions to LEO and especially the ISS, I can see it happening.
The heat shield for the Dragon base is around 600 kg, IIRC. A second stage might need more down the sides, but it is also a lot less dense than a Dragon capsule, which should mean less heat load because it decelerates already higher up in the atmosphere. So, I think it is safe to assume a TPS weight of less than 1000 kg. Then we we need to add some recover equipment depending on how they are going to land it (water drop with parachutes, propulsive landing, etc). I think they have routinely done deorbit burns with the second stage for LEO missions already. So that fuel is probably already part of the standard mission planning.
Anyone got some numbers for that one?
I think the real big question marks are the numbers for the recovery equipment and fuel and there is a lot of possible variations for how that could play out.

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