Author Topic: F9 - S2 reusable modification as evolution steps to BFS(ITS)  (Read 89615 times)

Offline livingjw

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A general rule of thumb for aircraft it 4% of landing or 3% of max takeoff weight. Which ever is heavier. Also, approximately 45% of that weight is wheels, tires and brakes. These estimates are for high strength steel and aluminum. My estimate based on use of carbon composite, would be no more than 2% of max landing weight, or there abouts. Mason has a very good report on estimating landing gear based on applied stresses. Enough math to get a fair estimate of the legs and oleos. Just google "Mason landing gear design". Additional weight for upper connections points around the tank should also be included. This might drive it above 2%. Did I see a max descent rate at touch down specified by Elon earlier. I think we have enough info to make an informed analysis.

John
« Last Edit: 08/30/2017 01:33 am by livingjw »

Offline john smith 19

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This is about US reusability and the long posts above arguing that US recovery is so difficult because mass mass mass.
Mass (or rather mass estimating the parts) is only part of the problem.

As an earlier poster noted the first stage landing gear weighs about the same as a Tesla.

Would you have the US come in with the same landing speed, and therefor likely need a landing gear of proportionate size? Or slower? Or faster?

But (per the title of this thread) you're going for a lifting entry, so high angle of attack while structure is bottom heavy. Note the term "drag" fins. Good at slowing down, not so good at generating lift.  :( And AFAIK SX have never put any on a US to test the atmospheric conditions at those altitudes, which will be significantly different.

This is both a formidable structures and a formidable control problem. Mass is a significant part of it but it's a long way from being all of the problems they will face. Keeping most of the stresses along the booster made booster recovery a lot easier. Once they go off axis that stage will get a lot heavier.
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Offline meekGee

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This is about US reusability and the long posts above arguing that US recovery is so difficult because mass mass mass.
Mass (or rather mass estimating the parts) is only part of the problem.

As an earlier poster noted the first stage landing gear weighs about the same as a Tesla.

Would you have the US come in with the same landing speed, and therefor likely need a landing gear of proportionate size? Or slower? Or faster?

But (per the title of this thread) you're going for a lifting entry, so high angle of attack while structure is bottom heavy. Note the term "drag" fins. Good at slowing down, not so good at generating lift.  :( And AFAIK SX have never put any on a US to test the atmospheric conditions at those altitudes, which will be significantly different.

This is both a formidable structures and a formidable control problem. Mass is a significant part of it but it's a long way from being all of the problems they will face. Keeping most of the stresses along the booster made booster recovery a lot easier. Once they go off axis that stage will get a lot heavier.
No it's not, as noted above.

Even though mass penalty for a second stage is 1:1, for LEO missions the payload is so much heavier than the US that even a doubling of second stage empty mass (which is very pessimistic) is perfectly acceptable.

GEO is a different matter, but US reusability is for high flight rate LEO​ missions.
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Offline john smith 19

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No it's not, as noted above.

Even though mass penalty for a second stage is 1:1, for LEO missions the payload is so much heavier than the US that even a doubling of second stage empty mass (which is very pessimistic) is perfectly acceptable.

GEO is a different matter, but US reusability is for high flight rate LEO​ missions.
I think I see where your logic is going wrong. You seem to be thinking that the recovery mass will scale with the size of the stage.

It doesn't.  :(

In order to keep the numbers simple I'll assume the booster separates at 1/2 orbital speed and 1/4 orbital altitude.  Both stages have both kinetic and potential energy. To land both stages you have to lower both their KE and PE to zero. I'll think in terms of a single Kg of mass.

So
S1 KE + PE = 1/2 m v(staging)^2 + mg x height_of_staging
S2 KE + PE = 1/2 m v(orbital)^2 + mg x  orbital_height

But if v(orbital) = 2x v(staging) and  orbital_height is 4x  height_of_staging
That means total energy is 4x per unit mass from orbit to that of 1/2 staging

That either means a) The US TPS has to have an erosion rate 4x less than the the booster TPS or 4x more of it will be ablated away.  :( Yes I think I'm beginning to see why Musk called it "uneconomic"

And that's with all things being equal.

Except they are not  :( because
a) The Merlin Vac nozzle is much bigger than it's first stage counterparts and
b) This assumes the trajectory that worked to lose the energy for the booster (but longer) can work for the US.  There's a lot of devils in those details.
c) Per the title of this thread people don't want to reenter like the first stage, they want to try out ways to do so like ITS IE a side on lifting entry

But if SX could do that wouldn't they have already fitted it to their boosters?

BTW regarding PICAX GW Johnson makes some interesting comments on PICAX, saying SX is in their 3rd generation and their goal is lowering the price of it.



IDK Wheather SX can make a PICAX 4.0 that's 4x better than the current grade. I guess it depends how much "stretch" PICAX has left in its thermal properties.
Science can't be rushed, and TPS development involves a lot of science.

So no TPS mass does not scale down with stage size. Couple that with having to protect a bigger nozzle and wanting to try an ITS landing instead (with it's much more complex control issues) and this is not going to happen anytime soon.
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Offline meekGee

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No it's not, as noted above.

Even though mass penalty for a second stage is 1:1, for LEO missions the payload is so much heavier than the US that even a doubling of second stage empty mass (which is very pessimistic) is perfectly acceptable.

GEO is a different matter, but US reusability is for high flight rate LEO​ missions.
I think I see where your logic is going wrong. You seem to be thinking that the recovery mass will scale with the size of the stage.

It doesn't.  :(

In order to keep the numbers simple I'll assume the booster separates at 1/2 orbital speed and 1/4 orbital altitude.  Both stages have both kinetic and potential energy. To land both stages you have to lower both their KE and PE to zero. I'll think in terms of a single Kg of mass.

So
S1 KE + PE = 1/2 m v(staging)^2 + mg x height_of_staging
S2 KE + PE = 1/2 m v(orbital)^2 + mg x  orbital_height

But if v(orbital) = 2x v(staging) and  orbital_height is 4x  height_of_staging
That means total energy is 4x per unit mass from orbit to that of 1/2 staging

That either means a) The US TPS has to have an erosion rate 4x less than the the booster TPS or 4x more of it will be ablated away.  :( Yes I think I'm beginning to see why Musk called it "uneconomic"

And that's with all things being equal.

Except they are not  :( because
a) The Merlin Vac nozzle is much bigger than it's first stage counterparts and
b) This assumes the trajectory that worked to lose the energy for the booster (but longer) can work for the US.  There's a lot of devils in those details.
c) Per the title of this thread people don't want to reenter like the first stage, they want to try out ways to do so like ITS IE a side on lifting entry

But if SX could do that wouldn't they have already fitted it to their boosters?

BTW regarding PICAX GW Johnson makes some interesting comments on PICAX, saying SX is in their 3rd generation and their goal is lowering the price of it.



IDK Wheather SX can make a PICAX 4.0 that's 4x better than the current grade. I guess it depends how much "stretch" PICAX has left in its thermal properties.
Science can't be rushed, and TPS development involves a lot of science.

So no TPS mass does not scale down with stage size. Couple that with having to protect a bigger nozzle and wanting to try an ITS landing instead (with it's much more complex control issues) and this is not going to happen anytime soon.
Nobody suggest that US recovery would be similar to boosters recovery.

If anything, it is similar to Dragon recovery.

You need a heat shield, some structural reinforcement, some propellant, something like grid fins, and a parachute.

If it were up to me, I'd eject the radiative nozzle.

These components do not add up to near the empty weight of the stage, and so are fine for LEO recovery.





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Offline Lars-J

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Nobody suggest that US recovery would be similar to boosters recovery.

If anything, it is similar to Dragon recovery.

You need a heat shield, some structural reinforcement, some propellant, something like grid fins, and a parachute.

If it were up to me, I'd eject the radiative nozzle.

These components do not add up to near the empty weight of the stage, and so are fine for LEO recovery.

Dropping the nozzle would aid simplify recovery but would negatively impact the refurbishment for re-use. Perhaps acceptable for some test flights, but it won't get you a system that closely mirrors the final destination. We do already know that the Vacuum Raptor will be actively cooled all the way to the edge of the vacuum nozzle (3m diameter), which suggests that they do not want to dump any nozzles in the long term. SpaceX wants to expend as few pieces as possible.

Offline meekGee

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Nobody suggest that US recovery would be similar to boosters recovery.

If anything, it is similar to Dragon recovery.

You need a heat shield, some structural reinforcement, some propellant, something like grid fins, and a parachute.

If it were up to me, I'd eject the radiative nozzle.

These components do not add up to near the empty weight of the stage, and so are fine for LEO recovery.

Dropping the nozzle would aid simplify recovery but would negatively impact the refurbishment for re-use. Perhaps acceptable for some test flights, but it won't get you a system that closely mirrors the final destination. We do already know that the Vacuum Raptor will be actively cooled all the way to the edge of the vacuum nozzle (3m diameter), which suggests that they do not want to dump any nozzles in the long term. SpaceX wants to expend as few pieces as possible.
Of all pieces, the radiative nozzle is the easiest to integrate, since it  is just a mechanical connection.

SpaceX is practical.  At some point it becomes a matter of cost.

I'll be thrilled if they can recover the nozzle too, but if that's getting in the way of recovering the stage, then so be it.

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Offline john smith 19

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Nobody suggest that US recovery would be similar to boosters recovery.
Engineers really like to leverage lessons learned, especially lessons that have cost a lot of time, pain and money to acquire. The video SX released suggested in fact that US recovery was going to be quite a bit like booster recovery. Since both start from roughly the same attitude, and have to end up in the same attitude (at least that was the plan) and are constructed in a very similar way a similar path seems like a pretty good starting point.
Quote from: meekGee 
If anything, it is similar to Dragon recovery.
You're going to have explain that one.

Dragon is a relatively small shape with quite a high density and most (all) of it sits inside the shock wave generated by the heat shield, with the most intense heating likely on a small area of the  lip between the windward face and the back side, the most forward part of the shape.

The US aspect ratio puts a lot of the rest of the structure in the airstream. "Hide in the shadow of the heat shield" is not a strategy that works with the shape of a structure the US has.

So could you explain what you mean by the US reentry being like a Dragon capsule?

Again this is entirely a consequence of trying to make something that's good at its job do something no other version of it has ever had to do, and has never been considered to have to need to have done.

Quote from: meekGee 
You need a heat shield, some structural reinforcement, some propellant, something like grid fins, and a parachute.

Only if those grid fins are made of Molybdenum or something similar  :(.
Titanium looks very high temperature by Aluminum standards, but checking its forging temperature is instructive, as that's when alloys have softened to a point they can be worked readily.
Ti runs 790-1065c, but the Inconels (used to make the X-15 skin, although its highest speed run used a spray on ablative) start at 1150c. You might also look at what the NASP programme was considering for structural materials as hypersonic flight has sometimes been called continuous reentry.  :(

And I wouldn't presume the upper stage grid fins (or whatever) would be smaller. Bigger fins could lose more velocity at a lower temperature rise before you sink (picking up speed) into the thicker air. Trade reusable mass (fins) for mass you have to replace (TPS). So your "Dumbo" fins can still be in Titanium, but are heavier than your booster set.   

I presume you're talking about parachutes to side step the need for landing legs.
Assuming you're not asking for a high temperature 'chute (about 700g/m^2 Vs 70g/m^2 for the usual kind. Remember CF will burn in air above 500c, lower if you have thin strands of it. Another case where  "Carbons high temperature performance is amazing" claim, comes with a long list of caveats  :( )

But the stage now has to survive a very large "jerk load" which is a)Large and b)Negative. Also if you want to avoid the landing leg penalty you need the chutes on the side of the stage. Chutes control packages can be 1.5% of landed weight but you're going to need a load spreading harness as a point load on the side will likely rip the panel it's mounted on away from the stage, which will continue to fall.

So you're looking at a parachute package on the side of the stage going up, along with associated load spreading harness that can survive reentry.
 
Note it does not have to be a woven fabric, and if it is woven it could be of metal or inorganic fiber. It's not impossible, but AFAIK there is no knowledge base for this. This is usually the point at which I learn that someone has been using a woven metal parachute harness for some obscure task, without any drama, for decades  :) .

Quote from: meekGee 
If it were up to me, I'd eject the radiative nozzle.
Me too.  :) It's the easiest way to allow the engine to do a landing burn without severe flow separation and side steps needing any super draco thrusters and their associated propellants.

Historically US engines have been made with a thrust chamber joint below the throat so you can either run the nozzle "as is" at sea level, then add an extension for US use, or cut off even shorter and chose whatever length you want for optimum altitude. You have to add a "sea level" nozzle. That makes it easier to move.

The lightest weight joint option has been to do all the SL testing, then weld on the extension for any testing in some kind of altitude test chamber before installing on the stage.

But


Welding refractory alloy nozzle extensions is tricky and time consuming, hitting your turnaround time (this exercise is all about reuse, right?)

 So you want a "demountable" joint that that is light,reasonably cheap,  seals tight, doesn't separate when you don't want it to and separates cleanly when you want it to. This is the point a lot of old aerospace engineers start muttering "explosive bolts," but for reusability (and cost, and testability) SX don't like them (rightly so IMHO). Proving that separate / no separate condition is critical.

IOW that joint is conceptually simple, but a fairly major PITA to develop.  :(

But.

Haven't you just spent a chunk of time and effort to eliminate propulsive landing with parachutes, and which option matches the ITS flight profile, whose development this is meant to be supporting?

If you need both then you need to add both the parachute mass, load spreading harness mass and terminal maneuver propellant to the balance sheet.  :(

Quote from: meekGee 
These components do not add up to near the empty weight of the stage, and so are fine for LEO recovery.
Again it's not just the raw weight of the additional parts.

It's raw weight
and additional stiffening for recovery mode
and propellants for recovery mode (if needed)
and the complexity of the control problem to make all this work together
and the refurb time you'll need to get the stage ready again.

BTW given the US will erode 4x the level of PICAX (per unit mass of the US) the booster will need that suggests that if Shotwells comments about no more than 3 reuses of the booster then the same thickness of PICAX in the same locations would burn through in 1 attempted landing of the US.
If V5.0 of the booster can do 10 launches without serious refurb that suggests an US with the same level of protection could do 2 safely.

But that's just an amateur's view of the problem.

The professionals have much better tools. I've never heard of SX doing a test of US recovery in the 6 years since their announcement video for full F9/Dragon reuse, which is suggestive that it's quite hard.

TL;DR version. Scaling rules between booster and upper stages are non linear and counter intuitive.

On topic for this thread.

No, F9-S2 is a poor development environment for ITS.  :(

The best way to develop ITS is either with ITS, at full scale (to avoid scale issues) or a sub scale, full featured ITS (ITSy, ITSlite, ITS 0.9, whatever) with ITS engines, propellants, TPS and landing systems. 
« Last Edit: 08/31/2017 09:50 am by john smith 19 »
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline meekGee

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You've chosen to design US recovery like booster recovery, and then spend all this time showing how difficult it is.

But if you look at speed,  mass, aspect ratio - you see that the US is a lot closer to Dragon than it is to the bottom, that's all.
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Offline john smith 19

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But if you look at speed,  mass, aspect ratio - you see that the US is a lot closer to Dragon than it is to the bottom, that's all.
Explain how it is and how you plan to make US recovery work. and how (given the title of the thread) this relates to developing ITS better/faster/cheaper.

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Offline spacenut

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I'm not a rocket scientist, but I see two choices. 

One, the weight of heat shielding materials and side re-entry, then go vertical for landing or parachute landing.  Some fuel needed to slow to re-enter, and then to land or use chutes.

Two, the weight of extra fuel to slow down like a booster, and enough fuel to land like a booster. 

Which is the lightest of the two, heat shielding materials or extra fuel. 

If it is going to be a test bed for ITSy, then a stretched stage, extra fuel and some side entry heat shielding.  Then, only FH could launch the stage.  F9 maybe not so much unless it is a widened stage to say 5m to match the fairing. 

Offline john smith 19

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I'm not really sure what this thread is for.  :(

Looking at the OP the premise seems quite far fetched to begin with and I'm not sure anyone has explained how changes to the F9 S2, or tests on it, will help development of ITS.

The F9 S2 is so different to the ITS as it has been described (before any changes in the revised versions Musk is going to present), in materials, engines, propellants and shear size, that it's hard to see how those lessons would transfer over.

It's only real benefits to doing this are it's flying now and its much smaller, so easier to modify. OTOH following SX usual rule that they do their testing as part of existing flight missions any changes would have to be agreed with their customers, and the further the design diverges from the standard F9 S2 the riskier this gets.

It sounds implausible because it is implausible.
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Offline Req

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But if you look at speed,  mass, aspect ratio - you see that the US is a lot closer to Dragon than it is to the bottom, that's all.
Explain how it is and how you plan to make US recovery work. and how (given the title of the thread) this relates to developing ITS better/faster/cheaper.

https://forum.nasaspaceflight.com/index.php?topic=43374.msg1707753#msg1707753

As far as how it may apply to the ITS, the re-entry attitude and TPS strategies would likely be similar.
« Last Edit: 08/31/2017 07:28 pm by Req »

Offline john smith 19

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https://forum.nasaspaceflight.com/index.php?topic=43374.msg1707753#msg1707753

As far as how it may apply to the ITS, the re-entry attitude and TPS strategies would likely be similar.
So you think the US can be adapted to be recovered like the fairing, using parachute?

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Offline john smith 19

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You've chosen to design US recovery like booster recovery, and then spend all this time showing how difficult it is.

But if you look at speed,  mass, aspect ratio - you see that the US is a lot closer to Dragon than it is to the bottom, that's all.
"is a lot closer" means what exactly?

Key elements for US recovery to work will include what's the Cg and Cp, which will decide how the stage naturally falls and for TPS how much of the stage is in the leeward side.
Cp and Cg are completely different for US and Dragon.
In the case of Dragon most of the capsule is narrower than the heat shield . In the case of the US it isn't.

Again, isn't this meant to be helping the design of the ITS, not Dragon?
If ITS is meant to function as a lifting body then S2 will as well.

Try to visualize the forces on the stage during during firing, then try and visualize how those forces change (in both direction, magnitude and sign) on the stage when whatever recovery idea you think will work is put into effect.

Anything that does not match the pattern of the operating stage means the stage will need modifying.
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Offline oiorionsbelt

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Will super sonic retro propulsion play a part in second stage re entry?

Offline guckyfan

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Anything that does not match the pattern of the operating stage means the stage will need modifying.

On the last RTLS we have seen the stage with a significant angle of attack. That was a first stage, long and fragile. A second stage is much shorter. It can reenter with a very small angle of attack initially then gradually increasing angle when speed goes down. It will need flaps at the engine end to provide drag, steering and protection for the engine. Just like ITS and like IXV.


Offline john smith 19

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Anything that does not match the pattern of the operating stage means the stage will need modifying.

On the last RTLS we have seen the stage with a significant angle of attack. That was a first stage, long and fragile. A second stage is much shorter. It can reenter with a very small angle of attack initially then gradually increasing angle when speed goes down. It will need flaps at the engine end to provide drag, steering and protection for the engine. Just like ITS and like IXV.


Looking up the IXV shows it's empty weight is 480Kg and its loaded mass 1900

https://en.wikipedia.org/wiki/Intermediate_eXperimental_Vehicle

IOW it's structure is 25% of it's loaded weight and has no massive engine block at the back.

You really don't see how this is quite different to an actual stage, with a mass fraction that Musk says is nearer 3%?

What makes US recovery from LEO so difficult is not any one thing.

It's that all of those things have to come together at the same time.  :(

Designing ITS is a good idea.

Making the F9 US recoverable is a good idea.

Using the process of trying to turn the F9 US into a reusable stage to drive ITS design is a bad idea.  :( They have such different scales and such vastly different goals that there is very little that the US can teach the ITS design and turning the US into a model ITS is likely to break the design, either in mass growth or in control authority or being impossible to refurbish, or of course all three.

It's the difference between considering options based on a disciplined use of imagination and total fantasy.  :(
« Last Edit: 09/01/2017 10:48 am by john smith 19 »
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Offline guckyfan

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You really don't see how this is quite different to an actual stage, with a mass fraction that Musk says is nearer 3%?

You really don't see that I just point out a basic principle? It is the same basic principle as Elon Musk described for ITS.

Offline rsdavis9


On the last RTLS we have seen the stage with a significant angle of attack. That was a first stage, long and fragile. A second stage is much shorter. It can reenter with a very small angle of attack initially then gradually increasing angle when speed goes down. It will need flaps at the engine end to provide drag, steering and protection for the engine. Just like ITS and like IXV.

Looking up the IXV shows it's empty weight is 480Kg and its loaded mass 1900

https://en.wikipedia.org/wiki/Intermediate_eXperimental_Vehicle


So offtopic:
The article says
Quote
first lifting body to make reentry from orbital speed.
Isn't the space shuttle a lifting body with reentry from orbital speed?

EDIT: well to answer my own question...
I guess the space shuttle had wings and the IXV does not.


« Last Edit: 09/01/2017 02:57 pm by rsdavis9 »
With ELV best efficiency was the paradigm. The new paradigm is reusable, good enough, and commonality of design.
Same engines. Design once. Same vehicle. Design once. Reusable. Build once.

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