Author Topic: F9 - S2 reusable modification as evolution steps to BFS(ITS)  (Read 27136 times)

Offline raketa

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1/I think that Elon's decision to reuse S2 is not primarily driven by cost reduction
2/F9-S2 could serve as a small scale platform to finalize BFS shape, heat protection and flying/landing algorithm
3/F9-S2 could through small modification take approximate shape of BFS(ITS)
4/F9-S2 Heavy will be probably in dimension scale ~25% of BFS(ITS) and have also additional lifting capacity, to carry modification to orbit and still deliver a payload to LEO.
5/F9H with S2R could deliver cargo for paying customers and at the same time learn lessons for the development of BFS.
6/List of modification, that could be gradually implemented step by step and base on the previous flight experience:

Phase 1-Heat protection and small wings for direction purposes:
a/How much heat protection is necessary to save the tank from overheating:
if I use Space shuttle as a reference, heat protection will weight ~1ton
b/Small wings to direct S2 shuttle like approach through the atmosphere
weight ~ 1ton
c/Additional N2 for direction thruster
weight ~ several hundred kg

Phase 2 - Modified wings to optimized flight dynamics
Evolve wings shaping S2 to be even more similar to BFS
additional weight ~1-2 tons

Phase 3 -Add landing Thruster
Probably use Draco engine to softly land S2.
additional weight ~ 1 ton
I think S2 could gradually evolve to be reusable and give SpaceX great lessons for BFS finalization:
Phase 1: -survive heat
Phase 2: -improve cross landing and aerodynamic braking  characteristic
Phase3: -test landing approach and algorithm
Sounds like the modification weight could be ~5-7 ton, that for F9H could be reasonable.
Here is picture of mine idea how F9-S2 could be modified F9-S2-Reusable

Online Robotbeat

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Gotta wonder if the Falcon Heavy Demi flight upper stage will look something like that. Doubt it's practical operationally, but should be useful as a tech demo for BFS.
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Offline cambrianera

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For what is worth... from an old thread.
For recovery gear, a 20 mm thick LI-900 heatshield would have a mass of less than 200 kg.
And landing speed should be low enough to use landing skids efficiently (wingload is less than 80 kg/square meter).
Oh to be young again. . .

Offline Peter.Colin

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I also suspect the reuse of S2 is not primarily cost driven but for ITS development purposes.
What needs to be tested first are the hardest parts,
Raptor engine, landing cradle, refueling in orbit.
Heat shield also but there is probably less risk in it not working as envisioned.


Phase 0) Titanium grid fins on maybe a bit longer cylindrical S2

This configuration might not need a big heat shield to re-enter Earths atmosphere, just a lot of fuel, and much lower payload.

Phase 1) Landing in a landing cradle, might be possible with simple S2-gridfin, steering thrusters configuration but likely to fail several times (a Hail Mary attempt). It's better to use a S2 than an already re-usable S1 for these otherwise expensive landing cradle attempts. Grasshopper attempts will probably be done first.

If the cradle landing doesn't work, landing legs are needed, or super Draco engines for more a precise landing in the cradle

Phase 2) Raptor engine(s) for S2

Phase 3) Methalox S2 refueling in orbit

Refueling in orbit by another "tanker S2", might make heavier payloads and enough fuel for landing possible with this reusable S2 version.


Phase 4) Change the body to ITS Spaceship form

Landing Legs and Super Draco will make it work with more certainty but are non-essential for ITS development
« Last Edit: 07/16/2017 07:39 PM by Peter.Colin »

Online GORDAP

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I also suspect the reuse of S2 is not primarily cost driven but for ITS development purposes.
What needs to be tested first are the hardest parts,
Raptor engine, landing cradle, refueling in orbit.
Heat shield also but there is probably less risk in it not working as envisioned.


Phase 0) Titanium grid fins on maybe a bit longer cylindrical S2

This configuration might not need a big heat shield to re-enter Earths atmosphere, just a lot of fuel, and much lower payload.

Phase 1) Landing in a landing cradle, might be possible with simple S2-gridfin, steering thrusters configuration but likely to fail several times (a Hail Mary attempt). It's better to use a S2 than an already re-usable S1 for these otherwise expensive landing cradle attempts. Grasshopper attempts will probably be done first.

If the cradle landing doesn't work, landing legs are needed, or super Draco engines for more a precise landing in the cradle

Phase 2) Raptor engine(s) for S2

Phase 3) Methalox S2 refueling in orbit

Refueling in orbit by another "tanker S2", might make heavier payloads and enough fuel for landing possible with this reusable S2 version.


Phase 4) Change the body to ITS Spaceship form

Landing Legs and Super Draco will make it work with more certainty but are non-essential for ITS development


Landing in a landing cradle doesn't seem to be something SpaceX plans for the BFS (spacecraft) portion of the ITS system.

I think a more likely evolutionary scenario might be:

1) Develop a 'scaled' BFS as a reusable second stage for the FH.  This would have a 'captive' payload bay (no separate fairings) and would look (and land) like their published BFS renderings.  Diameter of 5.5-6 meters - whatever max would be feasible when launched with FH 1st stage(s).  From the start it would use a Raptor engine, composite tanks, and autogenous(sp?) pressurization.   This would primarily serve their constellation plans and high energy/mass GTO payloads.

2a) Develop a manned version of this craft (crew of 6-10).  Together with 2b (below), this would serve all manned LEO, cis-lunar, lunar landing, space tourism and Mars Exploration (versus colonization) purposes.  Musk's recent comments suggest he's about to make a concerted, public push to open up NASA to competitively bid all missions that SLS is currently planned for.  This craft would be SpaceX's offering.

2b) 'Tanker' version of this craft.  All 3 of these craft would have nearly identical outer mold lines.

3) Replace FH 1st stage(s) with single core 'Raptor 9'.

Offline envy887

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1) Develop a 'scaled' BFS as a reusable second stage for the FH.  This would have a 'captive' payload bay (no separate fairings) and would look (and land) like their published BFS renderings.  Diameter of 5.5-6 meters - whatever max would be feasible when launched with FH 1st stage(s).  From the start it would use a Raptor engine, composite tanks, and autogenous(sp?) pressurization.   This would primarily serve their constellation plans and high energy/mass GTO payloads.

2a) Develop a manned version of this craft (crew of 6-10).  Together with 2b (below), this would serve all manned LEO, cis-lunar, lunar landing, space tourism and Mars Exploration (versus colonization) purposes.  Musk's recent comments suggest he's about to make a concerted, public push to open up NASA to competitively bid all missions that SLS is currently planned for.  This craft would be SpaceX's offering.

2b) 'Tanker' version of this craft.  All 3 of these craft would have nearly identical outer mold lines.

3) Replace FH 1st stage(s) with single core 'Raptor 9'.

This evolutionary path has been discussed elsewhere on this forum, and the main drawbacks are the need to redesign the FH TEL and GSE, and difficulty in building a TEL and GSE that support a Raptor mini-BFS and the standard F9/FH upper stage.

The most obvious place to launch this vehicle would be 39A, but 39A is needed for NASA launches (especially crew) which require the current upper stage for the foreseeable future. The other obvious place to launch it is Boca Chica, which is great for Mars/Moon/GTO, but much less so for constellation launches due to launch azimuth constraints.

Launching anywhere else would require a new pad, so the only way this vehicle can support the initial constellation deployment in the 2019/2020 timeframe would be adding a crew tower to LC-40 and then modifying the 39A TEL/GSE to launch only FH+mini-BFS, while heavy NSS missions launch from Boca Chica and VAFB.

Offline Jim

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while heavy NSS missions launch from Boca Chica

Non started.  NSS missions only launch from the Cape and VAFB.

Offline rakaydos

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while heavy NSS missions launch from Boca Chica

Non started.  NSS missions only launch from the Cape and VAFB.
Historically, what other options were there, that NSS declined?

Offline Jim

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while heavy NSS missions launch from Boca Chica

Non started.  NSS missions only launch from the Cape and VAFB.
Historically, what other options were there, that NSS declined?

It is a top level requirement, there are no options.

Offline envy887

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while heavy NSS missions launch from Boca Chica

Non started.  NSS missions only launch from the Cape and VAFB.
Historically, what other options were there, that NSS declined?

It is a top level requirement, there are no options.

Has anyone offered a NSS launch from any other site and been refused?

Offline Ionmars

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If this approach were used to establish a platform for testing BFS, there may be two basic questions. First, how much mass will Falcon 9R block 5 boost to stage separation, and (2) how much volume could be built into a partially-fuelled, empty stage 2-to-orbit vehicle that resembles the BFS. This will tell us how closely this testbed BFS could resemble a final BFS with usable payload that could be launched by an intermediate BFR.
* Mars: a convenient service station for an asteroid-sized spaceship en-route to Ceres. *

Offline envy887

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If this approach were used to establish a platform for testing BFS, there may be two basic questions. First, how much mass will Falcon 9R block 5 boost to stage separation, and (2) how much volume could be built into a partially-fuelled, empty stage 2-to-orbit vehicle that resembles the BFS. This will tell us how closely this testbed BFS could resemble a final BFS with usable payload that could be launched by an intermediate BFR.

Why launch on F9 if FH is available and could launch it fully fueled?

Offline Jim

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Has anyone offered a NSS launch from any other site and been refused?

Nobody offers an NSS launch.  They respond to a solicitation that has launch site as a requirement.

Also, part of EELV certification is having operations located at the Cape or VAFB.

Offline envy887

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Has anyone offered a NSS launch from any other site and been refused?

Nobody offers an NSS launch.  They respond to a solicitation that has launch site as a requirement.

Also, part of EELV certification is having operations located at the Cape or VAFB.

What are the reasons for the requirement? Clearly not orbital mechanics.

Offline Jim

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What are the reasons for the requirement? Clearly not orbital mechanics.

The NSS has existing infrastructure at those launch sites to support its missions.

Offline envy887

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What are the reasons for the requirement? Clearly not orbital mechanics.

The NSS has existing infrastructure at those launch sites to support its missions.

Such as?

Offline Jim

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What are the reasons for the requirement? Clearly not orbital mechanics.

The NSS has existing infrastructure at those launch sites to support its missions.

Such as?

https://www.gao.gov/assets/660/652037.pdf

Offline Peter.Colin

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I also suspect the reuse of S2 is not primarily cost driven but for ITS development purposes.
What needs to be tested first are the hardest parts,
Raptor engine, landing cradle, refueling in orbit.
Heat shield also but there is probably less risk in it not working as envisioned.


Phase 0) Titanium grid fins on maybe a bit longer cylindrical S2

This configuration might not need a big heat shield to re-enter Earths atmosphere, just a lot of fuel, and much lower payload.

Phase 1) Landing in a landing cradle, might be possible with simple S2-gridfin, steering thrusters configuration but likely to fail several times (a Hail Mary attempt). It's better to use a S2 than an already re-usable S1 for these otherwise expensive landing cradle attempts. Grasshopper attempts will probably be done first.

If the cradle landing doesn't work, landing legs are needed, or super Draco engines for more a precise landing in the cradle

Phase 2) Raptor engine(s) for S2

Phase 3) Methalox S2 refueling in orbit

Refueling in orbit by another "tanker S2", might make heavier payloads and enough fuel for landing possible with this reusable S2 version.


Phase 4) Change the body to ITS Spaceship form

Landing Legs and Super Draco will make it work with more certainty but are non-essential for ITS development


Landing in a landing cradle doesn't seem to be something SpaceX plans for the BFS (spacecraft) portion of the ITS system.

I think a more likely evolutionary scenario might be:

1) Develop a 'scaled' BFS as a reusable second stage for the FH.  This would have a 'captive' payload bay (no separate fairings) and would look (and land) like their published BFS renderings.  Diameter of 5.5-6 meters - whatever max would be feasible when launched with FH 1st stage(s).  From the start it would use a Raptor engine, composite tanks, and autogenous(sp?) pressurization.   This would primarily serve their constellation plans and high energy/mass GTO payloads.

2a) Develop a manned version of this craft (crew of 6-10).  Together with 2b (below), this would serve all manned LEO, cis-lunar, lunar landing, space tourism and Mars Exploration (versus colonization) purposes.  Musk's recent comments suggest he's about to make a concerted, public push to open up NASA to competitively bid all missions that SLS is currently planned for.  This craft would be SpaceX's offering.

2b) 'Tanker' version of this craft.  All 3 of these craft would have nearly identical outer mold lines.

3) Replace FH 1st stage(s) with single core 'Raptor 9'.


Landing in a cradle is big part of the ITS Big Falcon Rocket so why not try to practice cradle landing using a lighter cylindrical S2? (The new positioning base thrusters can be less powerful)
For a small S2, a cradle landing saves weight, for a BFR I think their primary goal is not to save weight, but to save a transportation trip from landing pad to launching pad.. because it's so F... Big.

I agree that the desired smaller rocket will be a single core raptor first stage, and that there will be a satellite deploying BFS second stage on top of it.
Walking in developing steps backward from this end design to FH.
They might do such a mini BFS for FH, but it would surprise me a lot, if they began with the BFS form factor.

My guess is there will be cylindrical second stage steps in between.
I know it's kind of boring... but a "safer" development approach.

If they where to bid for an SLS alternative, a complete mini ITS (BFR/BFS) might be smarter than a complex FH/BFS.
Another SLS alternative is a reusable mini BFR to lift all the other companies spacecraft cheaper into orbit.
And of course also a mini BFS either part of the deal or not.


« Last Edit: 07/17/2017 09:14 PM by Peter.Colin »

Offline envy887

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What are the reasons for the requirement? Clearly not orbital mechanics.

The NSS has existing infrastructure at those launch sites to support its missions.

Such as?

https://www.gao.gov/assets/660/652037.pdf

Quote
Air Force officials indicated that existing sites at locations other than the Cape and Vandenberg were not comparably equipped for NSS launches; for example, they lack the necessary payload integration facilities.

Anything else? Building integration facilities doesn't seem to be beyond the realm of possibility, as SpaceX needs them anyway for commsats. What particular requirements do NSS payloads have for integration that a commercial bird wouldn't?

Offline raketa

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I also suspect the reuse of S2 is not primarily cost driven but for ITS development purposes.
What needs to be tested first are the hardest parts,
Raptor engine, landing cradle, refueling in orbit.
Heat shield also but there is probably less risk in it not working as envisioned.


Phase 0) Titanium grid fins on maybe a bit longer cylindrical S2

This configuration might not need a big heat shield to re-enter Earths atmosphere, just a lot of fuel, and much lower payload.

Phase 1) Landing in a landing cradle, might be possible with simple S2-gridfin, steering thrusters configuration but likely to fail several times (a Hail Mary attempt). It's better to use a S2 than an already re-usable S1 for these otherwise expensive landing cradle attempts. Grasshopper attempts will probably be done first.

If the cradle landing doesn't work, landing legs are needed, or super Draco engines for more a precise landing in the cradle

Phase 2) Raptor engine(s) for S2

Phase 3) Methalox S2 refueling in orbit

Refueling in orbit by another "tanker S2", might make heavier payloads and enough fuel for landing possible with this reusable S2 version.


Phase 4) Change the body to ITS Spaceship form

Landing Legs and Super Draco will make it work with more certainty but are non-essential for ITS development

1/Craddle landing
They don't need to practice cradle landing. I think BFR will have legs in first version land like F9 S1, until they will get confident and will try to land in cradle.

2/Raptor engine
I think definitely try  tested in S2 first, but Raptor is not available now.
My suggested modifications are very small they could try to implement right now.

Offline Ionmars

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If this approach were used to establish a platform for testing BFS, there may be two basic questions. First, how much mass will Falcon 9R block 5 boost to stage separation, and (2) how much volume could be built into a partially-fuelled, empty stage 2-to-orbit vehicle that resembles the BFS. This will tell us how closely this testbed BFS could resemble a final BFS with usable payload that could be launched by an intermediate BFR.

Why launch on F9 if FH is available and could launch it fully fueled?
F9 is less expensive for multiple test flights and is the base case for this thread. I don't know if it would have enough lift for the job. If not, then it will be FH or another approach.
* Mars: a convenient service station for an asteroid-sized spaceship en-route to Ceres. *

Offline Jim

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Anything else? Building integration facilities doesn't seem to be beyond the realm of possibility, as SpaceX needs them anyway for commsats. What particular requirements do NSS payloads have for integration that a commercial bird wouldn't?

just stop.    NSS has their requirements, if others want to play in their sandbox, then they have to play by NSS rules. 

Offline rakaydos

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Anything else? Building integration facilities doesn't seem to be beyond the realm of possibility, as SpaceX needs them anyway for commsats. What particular requirements do NSS payloads have for integration that a commercial bird wouldn't?

just stop.    NSS has their requirements, if others want to play in their sandbox, then they have to play by NSS rules.
There was a recent video of Elon telling the Arkansas govenor that one of the downsides of oldspace is that they will never say no to awful requirements.

Offline JamesH65

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Anything else? Building integration facilities doesn't seem to be beyond the realm of possibility, as SpaceX needs them anyway for commsats. What particular requirements do NSS payloads have for integration that a commercial bird wouldn't?

just stop.    NSS has their requirements, if others want to play in their sandbox, then they have to play by NSS rules.

Would it ever happen that no-one wants to play by those rules any more? Or charges so much money to stick to the rules  that even the NSS thinks they are taking the piss. Would/could the rules be changed?

Not saying they should do one or the other, just commenting that just maybe, over the next few years, rules could change.

Offline Jim

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There was a recent video of Elon telling the Arkansas govenor that one of the downsides of oldspace is that they will never say no to awful requirements.

And that is why it is unlikely that the NSS will use Falcon Heavy

Offline Jim

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Would it ever happen that no-one wants to play by those rules any more? Or charges so much money to stick to the rules  that even the NSS thinks they are taking the piss. Would/could the rules be changed?

Not saying they should do one or the other, just commenting that just maybe, over the next few years, rules could change.

Nope.  The security requirements and associated infrastructure is not going away, along with other infrastructure at DOD launch sites.  And that is what they are DOD launch sites.  Why should they have to move?  They are specifically set up to manage and handle the preparation and launch of NSS assets.


See EPF, TSF and SSF at the Cape.  OMRF, ASO 5m high bay and SSI PPF at VAFB.
« Last Edit: 07/18/2017 02:59 PM by Jim »

Online Basto

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Anything else? Building integration facilities doesn't seem to be beyond the realm of possibility, as SpaceX needs them anyway for commsats. What particular requirements do NSS payloads have for integration that a commercial bird wouldn't?

just stop.    NSS has their requirements, if others want to play in their sandbox, then they have to play by NSS rules.
There was a recent video of Elon telling the Arkansas govenor that one of the downsides of oldspace is that they will never say no to awful requirements.

To be clear this is not an "oldspace" issue. As Jim has said NSS launches from the cape because that is where their infrastructure is.

It would be like if a competitor to Six Flags opened up a theme park 3 states over in the middle of nowhere and expected you to go there because it's cheaper instead of the Six Flags that is down the road.

Offline rakaydos

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There was a recent video of Elon telling the Arkansas govenor that one of the downsides of oldspace is that they will never say no to awful requirements.

And that is why it is unlikely that the NSS will use Falcon Heavy
Out of context, that feels like a beer bet statement. Say, next 3 years, if there is no NSS flights on a falcon heavy, I send you a beer. If there is, I get one.

Offline meekGee

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If the NSS made the rules, the NSS can change the rules - if it wants to.


« Last Edit: 07/19/2017 05:12 AM by meekGee »
ABCD - Always Be Counting Down

Offline mikelepage

One thing I'm curious about - and SpaceX would have data on this now - is to what extent supersonic retro-propulsion reduces peak heating on the returning stage? and to what extent it reduces the need for shielding?

If the exhaust plume from the engine is creating a tear-drop-shaped protective interface around the returning stage, does this help to protect the stage? and/or create additional deceleration over & above the thrust itself in the same way a drogue chute does?

Basically I'm wondering to what extent the requirement for S2 shielding can be removed by just changing the reentry procedure (timing/throttling of reentry/landing burns).

Offline livingjw

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One thing I'm curious about - and SpaceX would have data on this now - is to what extent supersonic retro-propulsion reduces peak heating on the returning stage? and to what extent it reduces the need for shielding?

If the exhaust plume from the engine is creating a tear-drop-shaped protective interface around the returning stage, does this help to protect the stage? and/or create additional deceleration over & above the thrust itself in the same way a drogue chute does?

Basically I'm wondering to what extent the requirement for S2 shielding can be removed by just changing the reentry procedure (timing/throttling of reentry/landing burns).

Supersonic retro-propulsion will not be used, a heat shield will be. The only reason supersonic retro-propulsion is used for the booster is to keep it from exceeding about Mach 3.5. Much faster and the booster would get to hot. Reentry from orbit is 7 times faster than that and must use the atmosphere to slow down. Not enough propellant to do anything else. For landing it will either be parachutes or subsonic retro propulsion (the landing burn).

Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

John
« Last Edit: 07/19/2017 11:47 AM by livingjw »

Offline Req

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Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

John

Offline livingjw

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Thanks. That is what I was thinking of. After reading it, I saw no good reason to use retro-propulsion on Mars except for the actual landing. Aero-braking with low ballistic coefficient and some L/D will get you slowed down and ready for  the landing burn.

John
« Last Edit: 07/19/2017 04:18 PM by livingjw »

Offline Lars-J

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Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

No, no, no. People keep misunderstanding this. This observed effect was only at low throttle levels. It does not mean that braking is impossible with a center mounted engine. Surely you have observed the F9 braking burns?
« Last Edit: 07/19/2017 04:31 PM by Lars-J »

Offline envy887

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Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

No, no, no. People keep misunderstanding this. This observed effect was only at low throttle levels. It does not mean that braking is impossible with a center mounted engine. Surely you have observed the F9 braking burns?

Those do not rely to any significant extent on drag. Drag is most important in the hypersonic and high supersonic phases of entry.

Offline envy887

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One thing I'm curious about - and SpaceX would have data on this now - is to what extent supersonic retro-propulsion reduces peak heating on the returning stage? and to what extent it reduces the need for shielding?

If the exhaust plume from the engine is creating a tear-drop-shaped protective interface around the returning stage, does this help to protect the stage? and/or create additional deceleration over & above the thrust itself in the same way a drogue chute does?

Basically I'm wondering to what extent the requirement for S2 shielding can be removed by just changing the reentry procedure (timing/throttling of reentry/landing burns).

Stage 2 entry is at too high a velocity for the stage to do this. Firing the engine backwards decreases drag and decreases heating, but means it takes much longer to decelerate.

For example, at 5 g deceleration (50 m/s^2) it takes ~132 seconds to go from orbital velocity (~7790 m/s) to Mach 3.5 (~1190 m/s) where the heating rate is more acceptable. The MVac burns 103 kg/s at minimum throttle, so running it for 132 seconds requires ~13,700 kg of fuel... all of which has to be put in orbit in lieu of payload. That's a major issue when F9 can launch ~18,000 kg with booster reuse and most LEO payloads are 9,000 kg or more.

The problem with reduced drag is that it cannot dive deeper into the atmosphere and decelerate faster, at least not without significantly increased heating, which is what you're trying to avoid in the first place. A 20 g dive would cut the fuel requirements to ~3,400 kg, but greatly increase the TPS requirements. At that point, it's much simpler to use a fully orbital velocity capable TPS and skip the retro-propulsion at entry.

A orbital heatshield can also be easily uprated for entries from GTO, which are much more difficult with retro-propulsion.

Offline Peter.Colin

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One thing I'm curious about - and SpaceX would have data on this now - is to what extent supersonic retro-propulsion reduces peak heating on the returning stage? and to what extent it reduces the need for shielding?

If the exhaust plume from the engine is creating a tear-drop-shaped protective interface around the returning stage, does this help to protect the stage? and/or create additional deceleration over & above the thrust itself in the same way a drogue chute does?

Basically I'm wondering to what extent the requirement for S2 shielding can be removed by just changing the reentry procedure (timing/throttling of reentry/landing burns).

Supersonic retro-propulsion will not be used, a heat shield will be. The only reason supersonic retro-propulsion is used for the booster is to keep it from exceeding about Mach 3.5. Much faster and the booster would get to hot. Reentry from orbit is 7 times faster than that and must use the atmosphere to slow down. Not enough propellant to do anything else. For landing it will either be parachutes or subsonic retro propulsion (the landing burn).

Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

John

I think the S2 of a regular Falcon 9 could land without shielding, if the payload is replaced with an extra fuel tank, that can optionally be jetisonned.
It takes far less amount of fuel to decelerate from very high speed than to accelerate, because the S2 gets a lot lighter when burning fuel.

For the Falcon Heavy the second stage without shielding could be a little longer, and it would have enough fuel for a small usefull payload, a few deceleration burns and a landing burn.

If you would want to maximize the payload of a fully reusable Falcon Heavy, with a cylindrical S2.
Refueling in orbit definitely helps.
A Raptor engine also, as does landing in a landing cradle.

Heat shielding seems difficult for a cylindrical S2 because of the engine weight at the back end, that makes it unstable, you would have to ad steering fins etc.




« Last Edit: 07/19/2017 08:42 PM by Peter.Colin »

Offline Lars-J

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Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

No, no, no. People keep misunderstanding this. This observed effect was only at low throttle levels. It does not mean that braking is impossible with a center mounted engine. Surely you have observed the F9 braking burns?

Those do not rely to any significant extent on drag. Drag is most important in the hypersonic and high supersonic phases of entry.

Right. But the problem is that some people interpreted the paper on drag reduction using a central thruster as proof that retro propulsion at supersonic speeds is impossible with a central thruster. ("you'll never get thrust, only reduce drag!")
« Last Edit: 07/19/2017 06:06 PM by Lars-J »

Offline envy887

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Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

No, no, no. People keep misunderstanding this. This observed effect was only at low throttle levels. It does not mean that braking is impossible with a center mounted engine. Surely you have observed the F9 braking burns?

Those do not rely to any significant extent on drag. Drag is most important in the hypersonic and high supersonic phases of entry.

Right. But the problem is that some people interpreted the paper on drag reduction using a central thruster as proof that retro propulsion at supersonic speeds is impossible with a central thruster. ("you'll never get thrust, only reduce drag!")

Why wouldn't you get thrust? That defies conservation of momentum. I've never seen this claim before.

Offline livingjw

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Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

No, no, no. People keep misunderstanding this. This observed effect was only at low throttle levels. It does not mean that braking is impossible with a center mounted engine. Surely you have observed the F9 braking burns?

Those do not rely to any significant extent on drag. Drag is most important in the hypersonic and high supersonic phases of entry.

Right. But the problem is that some people interpreted the paper on drag reduction using a central thruster as proof that retro propulsion at supersonic speeds is impossible with a central thruster. ("you'll never get thrust, only reduce drag!")

Of course you can use retro-propulsion to slow down, but why would you want to? A heat shield is much much lighter way to slow down in an atmosphere at high speed. S2 will be using aerodynamic drag to slow down until just before landing.

John

Offline Lars-J

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Of course you can use retro-propulsion to slow down, but why would you want to? A heat shield is much much lighter way to slow down in an atmosphere at high speed. S2 will be using aerodynamic drag to slow down until just before landing.

Two reasons:
1. One Mars you will impact the ground at supersonic speed using only your heat shield
2. On Earth, the atmosphere is so thick at surface level that if you come in from the wrong angle, it is like hitting a wall. Too much G forces. (Which is why the F9 first stage does a braking burn)

Offline livingjw

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Of course you can use retro-propulsion to slow down, but why would you want to? A heat shield is much much lighter way to slow down in an atmosphere at high speed. S2 will be using aerodynamic drag to slow down until just before landing.

Two reasons:
1. One Mars you will impact the ground at supersonic speed using only your heat shield
2. On Earth, the atmosphere is so thick at surface level that if you come in from the wrong angle, it is like hitting a wall. Too much G forces. (Which is why the F9 first stage does a braking burn)

1. Aero-braking with very modest L/D and low ballistic coefficient will allow you to successfully enter the Martian atmosphere with relatively low max g's and slow to  about 1500 m/s. The final landing burn reduces this to zero. 
2. The F9 does a retro braking burn to limit its maximum reentry Mach to limit heating.

Offline wannamoonbase

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I'm really excited to hear EM confirm yesterday that they are going to be pursuing US reuse.

I'm looking forward to following the development and see what technology they come up with.
Excited to be finally into the first Falcon Heavy flow, we are getting so close!

Offline mikelepage

I'm really excited to hear EM confirm yesterday that they are going to be pursuing US reuse.

I'm looking forward to following the development and see what technology they come up with.

There was actually quite a lot to unpack in his statements yesterday.

He used to think that a "base heat shield plus landing legs" would be the best way to land on Mars, he no longer thinks this, so that's part of the reason why they nixed the landing legs/propulsive landing with Dragon 2.  Maybe that's referring specifically to landing legs that pop out of the heat shield and not ruling out F9 Booster style fold out landing legs, but still, it was interesting to hear him say it.

Very curious now to hear the ITS update talk.

Offline mikelepage



Btw.  I feel like the last dozen comments or so would be better informed if you watched from 15 minutes to 21 minutes of this video.


Offline livingjw

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Thank you. I watched it again. 5% reduction in landing fuel with more control over landing point, if done with peripheral rockets as apposed to center mounted engines. Is that your take on it as well?

John

Offline mikelepage

Thank you. I watched it again. 5% reduction in landing fuel with more control over landing point, if done with peripheral rockets as apposed to center mounted engines. Is that your take on it as well?

John

Yes. Except high level control over landing point is not just a "desirable", but absolutely critical for all human architectures, because of pre-deployed assets.
 
Studies have been done which show that a rocket firing from the center does not increase drag, but that drag can be increased if rockets are fired from around the perimeter. I cannot remember if heating was reduced. I will see if I can't dig up the study.

No, no, no. People keep misunderstanding this. This observed effect was only at low throttle levels. It does not mean that braking is impossible with a center mounted engine. Surely you have observed the F9 braking burns?

Those do not rely to any significant extent on drag. Drag is most important in the hypersonic and high supersonic phases of entry.

Right. But the problem is that some people interpreted the paper on drag reduction using a central thruster as proof that retro propulsion at supersonic speeds is impossible with a central thruster. ("you'll never get thrust, only reduce drag!")

Why wouldn't you get thrust? That defies conservation of momentum. I've never seen this claim before.

Hence my suggestion above Envy887, to watch the video (actually from 14 minutes - 21 minutes) - low level thrust displaces the bow shock from the vehicle, paradoxically reducing drag by more than the force of the thrust.

But most importantly, the video was showing that supersonic retropropulsion in specific cases can also increase the area of the bow shock front and double the effective thrust through drag when performed during this particular energy envelope.  This saves 5% landing fuel and increases cross range control to upwards of 100km - which will be virtually necessary if you want to land next to pre-deployed assets as will be the case in human landings.

Having said all that, this also answers my earlier question: The heat shield solves a completely different problem from the one supersonic retropropulsion solves, because peak heating occurs at much earlier in EDL, at mach 18 or so - when the vehicle is still moving at velocities far too fast to start retropropulsion (or deploy IADs for that matter).
 

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Of course you can use retro-propulsion to slow down, but why would you want to? A heat shield is much much lighter way to slow down in an atmosphere at high speed. S2 will be using aerodynamic drag to slow down until just before landing.

Two reasons:
1. One Mars you will impact the ground at supersonic speed using only your heat shield
2. On Earth, the atmosphere is so thick at surface level that if you come in from the wrong angle, it is like hitting a wall. Too much G forces. (Which is why the F9 first stage does a braking burn)

1. Aero-braking with very modest L/D and low ballistic coefficient will allow you to successfully enter the Martian atmosphere with relatively low max g's and slow to  about 1500 m/s. The final landing burn reduces this to zero. 
2. The F9 does a retro braking burn to limit its maximum reentry Mach to limit heating.
1) 1500m/s is definitely supersonic. So you need to still do Supersonic Retropropulsion.
2) Not just heating. Forces, too.
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Offline mikelepage

Getting back to the thread topic, I do wonder if, now that they've invested so much into improving the Falcon Heavy core stage, the way to test all the ITS technologies will take the form of a Falcon Heavy specific upper stage.

I'm picturing something spade shaped, which would sit across the top of all three FH boosters, since this would improve the aerodynamics/vibration environment during the boost phase.  If you put the heat shield on one face of the "spade" you get a much larger area across which you spread the energy of reentry, with recessed superdracos and legs to allow a belly landing. 

The idea would be to keep the Merlin Vac, RP1/Lox tankage exactly the same, but not simply have wings, but rather use the extra volume for superdraco tanks + landing apparatus, and potentially payload bays or RP1/Lox tanks to practice on-orbit refuelling.

PS: I made a crappy photoshop mockup to show it better.

Offline Req

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My impression, given the timing of Musk tweeting that they were again thinking about trying S2 recovery, was that it will just be TPS re-entry to a parafoil(drogues first?), then bouncy castle landing.

Beyond just the timing of the statement, this forum has been all up and down the topic of propulsive/winged/etc S2 recovery over the years, and it's hard to imagine any of those methods delivering enough payload to be used except for very infrequently, making it hard to justify development.  Superdracos in particular would be a real showstopper, they aren't going to add hypergolic handling to the flow for every rocket whether it has Dragon or not.  The problem has never been that it's not possible, just that Falcon 9 at it's size would have to eat up most of it's payload capability to do it.

Using the fairing recovery method, all you're taking away from the payload is the mass of the PICA-X and any flaps/aero surfaces that are needed, the weight of the propellant to do the de-orbit burn, the weight of systems to keep the stage alive(all of this so far is needed for ANY S2 recovery method), and the weight of the chutes.  It's a lot easier to imagine that closing both in payload penalty and development resources, especially given the already existing fairing recovery program.  It's also a method that could be kitted relatively easily, rather than having substantially different second stages for expendable and reusable.
« Last Edit: 07/29/2017 06:02 AM by Req »

Offline mikelepage

Getting back to the thread topic, I do wonder if, now that they've invested so much into improving the Falcon Heavy core stage, the way to test all the ITS technologies will take the form of a Falcon Heavy specific upper stage.

I'm picturing something spade shaped, which would sit across the top of all three FH boosters, since this would improve the aerodynamics/vibration environment during the boost phase.  If you put the heat shield on one face of the "spade" you get a much larger area across which you spread the energy of reentry, with recessed superdracos and legs to allow a belly landing. 

The idea would be to keep the Merlin Vac, RP1/Lox tankage exactly the same, but not simply have wings, but rather use the extra volume for superdraco tanks + landing apparatus, and potentially payload bays or RP1/Lox tanks to practice on-orbit refuelling.

PS: I made a crappy photoshop mockup to show it better.

I probably should have said it explicitly... The implication of this would be that they are planning single stick and triple stick variants of the ITS, like was shown in the "Falcon X" and "Falcon X heavy" diagrams that used to appear in SpaceX material.

The reason I think they might do this is because currently, the ITS "tanker" variant is supposed to do several trips in order to perform on orbit refuelling of the ITS "crew" variant that will go to Mars.  Perhaps instead, they could have a single stick ITS "crew", and a larger, triple stick ITS "tanker" (both reusable of course).

EDIT: the specific benefit is that the fuel which the tanker is taking to the ITS crew version is only required to stay cryogenic for the hours/days it would take to make a single trip up, refuel the crew craft, perform checks, then go for trans-Mars injection.

EDIT 2: So the practice exercise would be that an ordinary Falcon 9 takes up a multi ton probe - A Jupiter probe with Europa lander, or even a Uranus/Neptune probe - to LEO or higher, and then a Falcon Heavy launches soon afterwards with a "tanker" second stage.  This rendezvouses with the probe+upperstage, refuels it, which then allows a direct injection/fast transit to the outer planets for a fraction of the cost.
« Last Edit: 07/29/2017 07:43 AM by mikelepage »

Online Robotbeat

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Musk has basically ruled out triple core ITS.
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Offline mikelepage

Musk has basically ruled out triple core ITS.

Yeah I know.  But those statements were only considering the rocket science/physics of the problem.

The pros behind going back to a slightly smaller core diameter, triple stick ITS would have nothing to do with the rocket science, and everything to do with the logistics/business case.  Say that 1) the process of building FH has given them confidence that they've solved the unique issues with the triple stick config, and 2) they found the business case/logistics for reusable triple core with smaller core diameter (7-9m) is actually a cheaper/easier way to get a large mass to orbit, and 3) being able to launch all the fuel for the Mars vehicle in one go solves the cryogenic-fuel-transfer-in-orbit problem.

I dunno (I'm not set on this), I just think Elon's shown previously that he'll make trade-offs like that (favouring the less efficient but logistically simpler solution).

Offline Lars-J

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F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #53 on: 07/30/2017 06:58 AM »
I dunno (I'm not set on this), I just think Elon's shown previously that he'll make trade-offs like that (favouring the less efficient but logistically simpler solution).

Whether or not you are set on it doesn't matter. A multi-core ITS isn't happening.

ITS is difficult enough as it is. The complexity difference between F9 and FH should tell you the rest.
« Last Edit: 07/30/2017 06:14 PM by Lars-J »

Online MP99




I probably should have said it explicitly... The implication of this would be that they are planning single stick and triple stick variants of the ITS, like was shown in the "Falcon X" and "Falcon X heavy" diagrams that used to appear in SpaceX material.

The reason I think they might do this is because currently, the ITS "tanker" variant is supposed to do several trips in order to perform on orbit refuelling of the ITS "crew" variant that will go to Mars.  Perhaps instead, they could have a single stick ITS "crew", and a larger, triple stick ITS "tanker" (both reusable of course).

In addition to the other comments, a BFR centre stage would RTLS, not land on a barge like FH centre core. This would reduce the tanker payload to orbit.

Cheers, Martin

Online Ronsmytheiii

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Still think the simplest/easiest way for a second stage reuse is going to be an evolution of NASA's HIAD program, ULA is trying to use it for Vulcan 1st stage engine recovery so you dont need a massive redesign to "bolt it on"

Plus will eventually be useful for landing large payloads on Mars.

https://www.nasa.gov/directorates/spacetech/game_changing_development/HIAD/index.html

Edit screenshot from https://www.nasa.gov/sites/default/files/atoms/files/gcd_industryday_hiad.pdf
« Last Edit: 07/30/2017 08:38 AM by Ronsmytheiii »

Offline AncientU

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Still think the simplest/easiest way for a second stage reuse is going to be an evolution of NASA's HIAD program, ULA is trying to use it for Vulcan 1st stage engine recovery so you dont need a massive redesign to "bolt it on"

Plus will eventually be useful for landing large payloads on Mars.

https://www.nasa.gov/directorates/spacetech/game_changing_development/HIAD/index.html

Edit screenshot from https://www.nasa.gov/sites/default/files/atoms/files/gcd_industryday_hiad.pdf

Is there one stitch of evidence for SpaceX following this tech path, or just wishful thinking?
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Offline spacenut

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I would hope they would use a sub-scale Raptor vacuum in place of the Merlin Vacuum and maybe widen the upper stage.  This would get the return hardware on and still get the same sized payloads to orbit.  It could also be equipped with on orbit refueling to test landing use in space. 

The other way is to stretch the upper stage, but this may make it only for use with FH due to length and stresses on F9.  At least 3 cores would make it through the heavy atmosphere and keep each other from bending. 

Offline wannamoonbase

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I would hope they would use a sub-scale Raptor vacuum in place of the Merlin Vacuum and maybe widen the upper stage.  This would get the return hardware on and still get the same sized payloads to orbit.  It could also be equipped with on orbit refueling to test landing use in space. 

The other way is to stretch the upper stage, but this may make it only for use with FH due to length and stresses on F9.  At least 3 cores would make it through the heavy atmosphere and keep each other from bending. 

I think that would be exciting too, but many here think that is unlikely, that SpaceX will move onto the next vehicle generation.  Lots of changes to the pad equipment, new process and procedure and likely still 2 years from having a Raptor engine qualified.

US reuse, even with the M-Vac, which would be interesting from a capability and technology perspective.  But even that may eat up resources that delays the next generation of vehicle. 

Edit: I'm leaning toward them not doing US reuse.  The business case for $ payback with the F9/FH vehicles is harder to close. 
« Last Edit: 07/31/2017 03:30 PM by wannamoonbase »
Excited to be finally into the first Falcon Heavy flow, we are getting so close!

Offline titusou

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This is old SpaceX video (which not available on SpaceX's own channel) which show what they were thinking long time ago.


Having stage2 nose covered with heatshield and rentry as nose-first, as video suggested, well... 180deg flip back to nose-up at supersonic speed for landing burn would be quite challenging I guess.

Some other concept (not by SpaceX) has SuperDraco side-mounted (facing up) on stage2 for landing burn, which means Hydrazine engine next to LOX tank... well... :)

My personal favor would be a ITS upper stage liked design. Heatsheild one side, reentry as side-first, and then 90deg flip for landing burn. ITS upper stage is basically symmetric at launch. The extended heatshield under the engines, were merged with stage1 to make it symmetric at launch, asymmetric after separation.


Whatever design it will be, I hardly see any chance to have asymmetric shaped stage2, as it will generate complex force during liftoff. The symmetric shaped stage2 will make thing much less complicated.

Titus

Online guckyfan

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No problem with nose first reentry or only a small angle of attack. Then gradually switch to side entry when the speed is already reduced.

Online Robotbeat

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I'm pretty sure they'd do side entry like Shuttle or Delta Clipper.
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Offline spacenut

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Side entry would give them information on ITS or ITSY.  Top entry doesn't.  If it could flip after side entry and land engine down with Superdracos, that would be cool. 

Offline mikelepage

Question: do we know if grid fins can turn to an angle 90+ degrees from launch orientation?  Should be easy enough to do, right?

What I'm currently imagining for the "hail mary" attempt of the FH S2 is a second stage with side mounted heat shield/resistant coating(?), plus four grid fins: two grid fins are mounted either side of the heat shield at the top/front, and then two more grid fins mounted on the sides below these, next to the MVac, below/behind center of gravity.  No landing legs.  No super dracos.

Pacific ASDS hosts some kind of experimental cradle based on interstage connector.  S2 performs deorbit burn, then uses cold gas thrusters to orient stage side on to reentry.  Lower pair of grid fins deploy, turning right angles to stowed orientation - upper pair are (I think) counter productive at early stage when you want to use the side mounted heat shield and CoG is already so far towards the rear.

Flip to vertical landing orientation should be achievable without engine relighting if you have upper and lower pairs of grid fins, so well after peak heating (but before max Q I think), upper grid fin pair deploy and start flip motion.  Engine relight happens for landing burn, hopefully into the cradle mounted on ASDS. 

I believe the grid fins are only on the order of 50kg each, so not the biggest mass penalty.  The question in my mind is how heavy a side mounted heat shield is?

Offline alang

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Dumb question: does an inflatable heat shield have to cover the engine bell? Could it be done as a torus around the second stage engine?
Intuitively the answer is it wouldn't work because of the fragility of the second stage engine bell and the hypersonic re-entry, but having read the comments about the non-intuitive nature of aerodynamics and given the ubiquity of cheap computing power I wonder if anyone has actually modelled some of the more outlandish scenarios.
« Last Edit: 08/06/2017 10:49 AM by alang »

Offline tdperk

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It would be like if a competitor to Six Flags opened up a theme park 3 states over in the middle of nowhere and expected you to go there because it's cheaper instead of the Six Flags that is down the road.

How much cheaper?  Also, three of the big Western states or going to West CT From Attleboro, MA?

Offline First Mate Rummey

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Sorry for asking, but what do you mean with ITSY?
Thanks.

Sorry for asking, but what do you mean with ITSY?
Thanks.

ITSy is the nickname we've given to the subscale ITS. It's not official at all.

Offline Lar

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Sorry for asking, but what do you mean with ITSY?
Thanks.

ITSy is the nickname we've given to the subscale ITS. It's not official at all.

Keep using it and maybe it'll catch on!!!
"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 oldAtlas_Eguy

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As a version of and speculation if S2 is heavily modified then it may likely also see the use of carbon composite tanks. This would increase the capability to reduce the impacts made by additional reuse hardware on the payload capability of the US. Then if such major changes are made then also is a possibility of changing the stage out to use methalox with autogenous pressurization. That would also likely change the stage to also be a 5m diameter but the same length reducing impacts on towers and other items placement in the GSE in support of the US and payloads on top.

Offline Arb

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That would also likely change the stage to also be a 5m diameter but the same length reducing impacts on towers and other items placement in the GSE in support of the US and payloads on top.
That'd require reworking of the transporter erector to accommodate the increased diameter; how would that even work/look-like?

|
\
 |
 |
 |
 |
 ---
(transporter/erector not to scale)

It'd also be a pain when horizontal; with the first stage having to be somewhat higher off the ground than at present.
   ________|
__/


Interesting trade...

Edit: spelling.
« Last Edit: 08/11/2017 07:41 PM by Arb »

Offline GWH

I feel like this topic has been brought up lots before, but....

It looks to me like there could be space, however the cradle arms would need to move to the outside rather than fully under.

Two different angles shown in the photo below, the bottom one rotated in orientation to fit better.  Looks like 5m clear would fit the whole length if the arms weren't in place, and then note the shipping splice in the TEL that is located 1/3 of the way up the interstage.

Theoretically this upper section could simply be replaced, theoretically because I have no idea of the inner details.

There may have been a very recent image floating around where the TEL was erected and the upper section was removed, but I can't find where I saw that.
« Last Edit: 08/11/2017 12:09 AM by GWH »

Offline john smith 19

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I'm pretty sure they'd do side entry like Shuttle or Delta Clipper.
Only if you make the structure a) Winged or b) Lower aspect ratio.  :(

"Upper stage reuse" has been known to be possible since the first Shuttle reflight if you're prepared to design a vehicle to do it.

What has made it difficult is if you insist on doing it with a high aspect ratio US, which is what a conventional US is.

You are aware that the Shuttle reentry angle (when it lost most of its velocity) was rather nearer to 70degrees to the horizontal than horizontal?

Or the DC-X slightly nose down concept was inspired by the USAF's experience with warhead reentry?

This idea has a couple of inescapable facts connected with it.

1) Every unit of mass you put on the US comes off the payload. On boosters it's more like 13:1.  So you can add a (comparatively) huge amount of hardware to do recovery while sacrificing a (comparatively) small amount of payload.

2) You can make VTO LV's very light because they are very optimized to be strong in exactly 1 axis, just like a soda can  :( I've stacked 10 fully filled (but unpressurized, which is important as pressurization really helps) vertically. Try stacking with the 9 fully loaded cans on the side of a closed can and see what happens. Make sure you do it in a bowl. It's going to get messy. :(

We know that a layer of paint is enough to insulate a LOX tank through 1st stage recovery engine on, but I don't think that's going to hack it for a side entry (let along not exploding due to flash boiling), do you?

So (assuming you've beefed up the whole tank structure to handle substantial side loads) how will you handle the heat, given you have to retain propellant for landing?

Obvious options I can see are
a)Minimal mass by putting TPS on the "windward" side of the tank.
b)Even layer of TPS on the sides of the whole stage
c)Thinner layer of TPS on whole stage and slow spin the stage on landing ("Rotisserie" mode as they called it for Apollo and Saturn EDS).

All of these offer exciting opportunities to advance the science of control systems quite a lot, with the possible exception of b), Which is the simplest (and heaviest) option.  :)

Fortunately the US Army has done quite a lot of work on the guidance issues associated with pairs of spinning tanks in tandem during their development of binary nerve gas shells (and some on bombs), as the consequences of a stray shell were quite serious.
Obviously the tanks are rather smaller (and faster spinning) but AFAIK the mathematics still applies wheather it be 50Kg or 5000Kg of fluid and 1 or 1000RPM. Simply insert appropriate operating conditions and fluid properties.

But remember
a)All require the interior structure and/or the tank pressures to be increased, since this is completely unnecessary for current expendable US practice
and
b) the "exchange rate" for US is 1:1.
Every lb/Kg/banana of mass you add comes off the payload reserve (which is no longer the payload reserve, and you'd be misleading people if you called it that) or off the available payload to orbit.

With the exception of tank pressurization gas (which could in principle be vented, if you could retain the propellant) every unit of mass that goes to orbit on the stage (except the payload) has to come back, and be decelerated in the final landing.  :(
« Last Edit: 08/28/2017 03:45 PM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline meekGee

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I'm pretty sure they'd do side entry like Shuttle or Delta Clipper.
Only if you make the structure a) Winged or b) Lower aspect ratio.  :(

"Upper stage reuse" has been known to be possible since the first Shuttle reflight if you're prepared to design a vehicle to do it.

What has made it difficult is if you insist on doing it with a high aspect ratio US, which is what a conventional US is.

You are aware that the Shuttle reentry angle (when it lost most of its velocity) was rather nearer to 70degrees to the horizontal than horizontal?

Or the DC-X slightly nose down concept was inspired by the USAF's experience with warhead reentry?

This idea has a couple of inescapable facts connected with it.

1) Every unit of mass you put on the US comes off the payload. On boosters it's more like 13:1.  So you can add a (comparatively) huge amount of hardware to do recovery while sacrificing a (comparatively) small amount of payload.

2) You can make VTO LV's very light because they are very optimized to be strong in exactly 1 axis, just like a soda can  :( I've stacked 10 fully filled (but unpressurized, which is important as pressurization really helps) vertically. Try stacking with the 9 fully loaded cans on the side of a closed can and see what happens. Make sure you do it in a bowl. It's going to get messy. :(

We know that a layer of paint is enough to insulate a LOX tank through 1st stage recovery engine on, but I don't think that's going to hack it for a side entry (let along not exploding due to flash boiling), do you?

So (assuming you've beefed up the whole tank structure to handle substantial side loads) how will you handle the heat, given you have to retain propellant for landing?

Obvious options I can see are
a)Minimal mass by putting TPS on the "windward" side of the tank.
b)Even layer of TPS on the sides of the whole stage
c)Thinner layer of TPS on whole stage and slow spin the stage on landing ("Rotisserie" mode as they called it for Apollo and Saturn EDS).

All of these offer exciting opportunities to advance the science of control systems quite a lot, with the possible exception of b), Which is the simplest (and heaviest) option.  :)

Fortunately the US Army has done quite a lot of work on the guidance issues associated with pairs of spinning tanks in tandem during their development of binary nerve gas shells (and some on bombs), as the consequences of a stray shell were quite serious.
Obviously the tanks are rather smaller (and faster spinning) but AFAIK the mathematics still applies wheather it be 50Kg or 5000Kg of fluid and 1 or 1000RPM. Simply insert appropriate operating conditions and fluid properties.

But remember
a)All require the interior structure and/or the tank pressures to be increased, since this is completely unnecessary for current expendable US practice
and
b) the "exchange rate" for US is 1:1.
Every lb/Kg/banana of mass you add comes off the payload reserve (which is no longer the payload reserve, and you'd be misleading people if you called it that) or off the available payload to orbit.

With the exception of tank pressurization gas (which could in principle be vented, if you could retain the propellant) every unit of mass that goes to orbit on the stage (except the payload) has to come back, and be decelerated in the final landing.  :(
The 1:1 payload penalty is a red herring.

For LEO missions, a 5 ton US places a payload that's in excess of 20 tons for F9b5, not to mention FH.

So even if US reusability doubles the empty mass of the stage, a 1:1 penalty only drops the payload mass to 15 tons - hardly a show stopper... Not to mention that the mass penalty estimates are more like 3 tons.


Further, many LEO missions are volume limited.

There is absolutely nothing that is preventing SpaceX from building a recoverable upper stage for LEO missions.

All zat is rekvired is the VILL to do so!
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Online Robotbeat

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I'm pretty sure they'd do side entry like Shuttle or Delta Clipper.
Only if you make the structure a) Winged or b) Lower aspect ratio.  :(...
Why is your post full of frowny faces? They'll figure it out eventually. And in the meantime, we get explosions! Win, win! :D

Anyway, I'm sure the recovery-attempt upper stage will look all weird with funky aerosurfaces for Falcon Heavy's maiden launch. Like a bastardized ITS spaceship.

Turn that frown upside down! :D
« Last Edit: 08/29/2017 04:09 AM by Robotbeat »
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Offline john smith 19

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Further, many LEO missions are volume limited.

There is absolutely nothing that is preventing SpaceX from building a recoverable upper stage for LEO missions.
If you discount the fact landing gear on the first stage turned out considerably heavier than most early estimates expected (more like 7% of GTOW IIRC) and as I noted the interesting control systems problems of landing such a stage, especially if it comes in at anything more than a shallow angle to the horizontal IE 3-5 degrees, not 70+

Quote from: meekGee
All zat is rekvired is the VILL to do so!
Which was pretty much the claim made by stage mfg's and recovery advocates in the 1960's for first stage recovery.

All that was needed was the (government) money to demonstrate it.  :(

None of those concepts had grid fins at the top end of the stage because everyone thought they could do it with thrust vectoring the engines if any control was needed.

That tells me that every single one of those concepts was not worked out in sufficient detail to realize they were needed. 

IOW they were all fantasies:(

SX have set the bar for what it takes to do this IRL. No Powerpoints, no videos.  It took them 5 attempts to get a successful barge landing and right first time for a land landing, despite the Grasshopper test programme before they made any actual attempts. I'm quite sure data from the barge landings fed their first land attempt and made it successful.
Why is your post full of frowny faces? They'll figure it out eventually. And in the meantime, we get explosions! Win, win! :D
I've no doubt that if the problem can be solved at all the SX team is smart enough to do it.

As someone who's keen to see launch prices drop radically the fuller the reuse the better. In money terms I'm betting fairing reuse will have an earlier (positive) impact on their bottom line than US recovery but only US recovery gives the really big price drops.

Quote from: Robotbeat
Anyway, I'm sure the recovery-attempt upper stage will look all weird with funky aerosurfaces for Falcon Heavy's maiden launch. Like a bastardized ITS spaceship.

Turn that frown upside down! :D
I'll try to keep that in mind.

I prefer my development programmes explosion free. Great entertainment but they make potential customers very nervous.

It's not just the loud bangs.

Stage reuse is meant to improve reliability above the expendable level.

SG1962's idea that the actual FH US will be a stock FH US (which ideally will be just a stock F9 US, but may not be,  given how far the design of the FH core has diverged) and the recovery attempt will be tried by the payload makes a lot of sense to me.  It's a classic "Get the core tasks out of the way and see if we can make bonus features work" SX approach.

If it works, brilliant.  :) If it doesn't, FH has demonstrated flight readiness, possibly with full first stage core recovery (and at relatively low cost). A key task if they want to go after the bigger NSS payloads with higher orbits.
« Last Edit: 08/29/2017 08:01 AM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline meekGee

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7% GTOW - wouldn't that make the legs as heavy as the entire rocket?

Anyway, what legs?  Parachute, aim for near shore, mid-air recovery of an empty US. 

No need for landing engines or any of that.  Grid fins for control, and a heat shield.

That's why thinking a 5 tons penalty is excessive.

This here is the first time there's any real motivation to recover the upper stage.  (Because of LEO demand).

GEO recovery is just much harder, and the mass penalty becomes a showstopper.
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Offline envy887

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7% GTOW - wouldn't that make the legs as heavy as the entire rocket?
...

7% of LANDING weight, not takeoff weight. And mostly due to the large span required for stability. The upper stage is relatively short and fat.

Offline john smith 19

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7% GTOW - wouldn't that make the legs as heavy as the entire rocket?
...

7% of LANDING weight, not takeoff weight. And mostly due to the large span required for stability. The upper stage is relatively short and fat.
Thank you for reminding me. Note that the heuristic for HTO vehicles is 4% of landing weight, despite the fact the XB70 managed to do it for 2% as the Boeing page of XB70 shows.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline meekGee

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7% GTOW - wouldn't that make the legs as heavy as the entire rocket?
...

7% of LANDING weight, not takeoff weight. And mostly due to the large span required for stability. The upper stage is relatively short and fat.
Thank you for reminding me. Note that the heuristic for HTO vehicles is 4% of landing weight, despite the fact the XB70 managed to do it for 2% as the Boeing page of XB70 shows.
This is not about HTO vs VTO, or heuristics vs. real.... 

This is about US reusability and the long posts above arguing that US recovery is so difficult because mass mass mass.
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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.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

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.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

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 »
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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.
ABCD - Always Be Counting Down

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.

"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

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.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

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?

"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

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.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Online oiorionsbelt

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

Online 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 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Online 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 »
bob

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #100 on: 09/01/2017 03:25 PM »
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.
I'm not sure you're saying what you think you're saying.  :(

A launch stage is a launch stage and a spacecraft is a spacecraft. the F9-S2 is a stage. ITS is a spacecraft. It's the difference between the Shuttle External Tank and the Shuttle, and then trying to make the ET reusable and landable on a runway like the Shuttle.

There is nothing that is "basic" about this process, and most people would think anyone trying to do it was pretty stupid  :(

I think the SX design team have demonstrated they are anything but stupid.  I see as much likelihood of the F9 US "evolving" into ITS as I see a tortoise evolving into a sperm whale. But please continue to explain what you have in mind. If it's so simple it should be simple to explain. Perhaps with some numbers?

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.
The common meaning of "lifting body" is a vehicle that has no wings, and gets all its lift from the shape of the fuselage, whereas the common aircraft is (roughly) a tube with wings attached.
Since the tube normally carries something at pressure (fuel, LOX, passengers at near sea level pressure), and it's easier to make pressure vessels that are cylinders.

The SR71 and the shuttle were exceptions.
The SR71 outside was elliptical (housing various bays, generating some lift in the "strakes," but the actual pressure carrying structure was an inner cylinder.  Sort of a semi-lifting body.
The Shuttle was essentially designed to equalize pressure with the outside during ascent and descent (only the nose cabin are was actually pressure tight), hence wings on basically a hump backed rectangular body (which may have also been shaped to allow analysis by the very limited CFD tools of the time  :( ).

Lifting bodies have flown from ICBM's and rockets as model tests ("Surviving the Heat Barrier," TA Heppenheimer) and Dream Chaser is a fully fledged LB (as well as being a human capable fully composite vehicle) but TBH I think they mean "First lifting body in Europe," which probably is completely true, but not quite so snappy.  :)
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline Space Ghost 1962

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #101 on: 09/01/2017 07:15 PM »
Will try one post on this really pointless thread.

There's a conflation of too much in too useless packages.

Let's start with what they've got and why they got it.

A reusable booster with landing gear that launches from a R/ELV pad and lands on a flat, otherwise unremarkable surface/barge.

They did this so that they could fly a vehicle ELV or RLV, and if the RLV failed on recovery, it wouldn't destroy too much (how many holes in the barge, lost recovery GSE?). Clearly the decision was for an interim vehicle CONOPs to function in the industry where ELV/RLV economics compete/coexist.

This placed the burden of the additional cost of RLV as excess performance to compensate for recovery (boostback,  landing, legs, grid fins ...).

The benefits of this were 1) booster reuse and 2) booster reuse operations experience.

Everyone focuses on 1) not 2). 2) is more valuable than 1).

Note that with the ITS reveal, BFS has legs like the booster,  BFR doesn't. Meaning that sometimes they use stuff differently (or not at all, as recently with Dragon 2). Pragmatics.

Now, lets look at the next business need. The big omission for cost/frequency in capturing more market share might be the expendable US, for if they do day to day reuse of the booster, sooner or later they'll run out of F9US's.

Sure, they could do a NG like larger scale vehicle in an ITSy. Would bring them closer to Mars goal soonest. But there's no Mars business yet. So the long term goal does not leverage the short term business advantage.

F9 "owns" them now. That's why FH is finally happening. Nothing about FH helps ITSy. If anything, it postpones ITSy (buys time to do it in having a partially reusable HLV). The "F9 business" then flies fewer ELV missions, to allow ELV to be "phased out".

(Note that NG is pictured as flying two methalox stages with a 5M-ish fairing, meaning existing payloads/business too.)

There is little likelihood that any of F9/FH/NG flying anything but current payloads for the foreseeable future.

A crewed Dragon 2 on a Block 5  is the closest to an unconventional  payload. Next is a possible lunar adventure with FH. That's it for now. Possibly longer term DIVH NSS like missions.

So what will benefit ITSy is not BFS like stuff with F9/FH, but "recoverable US operations". For that, you need ... to recover F9US.

How is F9/FH unlike ITSy? On the scale of ITSy, it must be 100% recoverable, while F9/FH must allow for ELV as part of its scaling.

The cornerstone of the entire Falcon launch provider business is a refined F9US. Any changes impact all missions. And, it is the single largest remaining cost to every launch.

From a competitive perspective, one can initially compete with a recoverable booster by economies of scale. (It may be a losing battle, but its still a battle.) However, if you then add a reusable, performant US ... then no economies of scale can save that.

F9US plus fairing is a system that works unlike BFS. You'll never approach similar design. Nor would you want to.

Falcon's kerolox economics come from matching an idealized propulsion/CONOPs model. Some improvements take that further, others fight the model.

Landing without legs a booster might improve the model, but not if there's bigger obstacles (threatening existing F9 operations, F9US limiting cost/frequency,...).

And most of all, they need to have the experience first of operating a fully reusable vehicle ... to be able to design/implement a more optimal, larger scale vehicle, done from the ground up as such.

Because then you get it right the first time.

Offline GWH


The benefits of this were 1) booster reuse and 2) booster reuse operations experience.

Everyone focuses on 1) not 2). 2) is more valuable than 1).

And most of all, they need to have the experience first of operating a fully reusable vehicle ... to be able to design/implement a more optimal, larger scale vehicle, done from the ground up as such.

Because then you get it right the first time.

Thank you.
Cannot count the number of time the importance of these points are missed when folks are discussing any type of development ahead of ITS.

Offline alang

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #103 on: 09/02/2017 04:05 PM »
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.

I know little and tend to ask similar questions but I don't understand why you think Falcon Heavy necessarily needs to have a stretched upper stage to improve a reuse demonstration, even if desirable. Is that a given?
If the upper stage on the FH core separates later than on F9 then perhaps there could be spare fuel for deorbiting to a speed that requires less thermal protection than from orbital speed. What I don't know is what reducing orbital speed by an arbitrary percentage will do to the need for thermal protection. Perhaps there is some critical velocity that reentry needs to be below in order to avoid the need for more exotic materials.
If we assume that the upper stage separates later with F9 then I suppose the core stage also needs more fuel for its reentry burn to reduce the additional thermal stress. This all assumes that the side boosters carry the core stage so it doesn't burn so much fuel before they separate.
Then there is the question of what the FH demo flight is meant to prove. If they can launch to orbit a mass simulator slightly more massive than the maximum mass of an F9 LEO payload with first stage recovery then how much second stage fuel could be left over when it reaches orbit? The SpaceX website only seems to provide a maximum payload for a non reusable FH launch.
« Last Edit: 09/02/2017 04:06 PM by alang »

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #104 on: 09/02/2017 11:30 PM »
What I don't know is what reducing orbital speed by an arbitrary percentage will do to the need for thermal protection. Perhaps there is some critical velocity that reentry needs to be below in order to avoid the need for more exotic materials.
You need a better understanding of what those terms mean for a start.

"orbital velocity" is the speed an object has to be moving around the Earth to balance the exact level of gravity at that altitude (the one you get when you plug the actual values for the earths radius and the object into the relevant formula).

But g rises as altitude falls, so once you slow down you will fall down.

So you can't be "reducing orbital speed by an arbitrary percentage" as you're not in orbit anymore, you're starting reentry.

Quote from: alang
If we assume that the upper stage separates later with F9 then I suppose the core stage also needs more fuel for its reentry burn to reduce the additional thermal stress. This all assumes that the side boosters carry the core stage so it doesn't burn so much fuel before they separate.
Or more TPS.

This problem is trickier than it looks. 

To come to land the stage has to lose basically all the energy it gained to get to orbit, which is a combination of kinetic and potential energy. Roughly speaking for S1 separation at 1/2 orbital velocity and 1/4 altitude (which keep the numbers simple) a unit of mass of the US has to lose 4x the energy a booster stage has to to land.

You can't decelerate completely using rocket thrust because you'll need as much propellant to slow down as you used to speed up. But even if you could bring it to a complete stop it will now fall 200Km straight down, picking up speed all the time.  Potential energy becomes kinetic energy and you still have to get rid of it.

This is on a "per mass" basis. IOW it doesn't matter if it's a small dense ball (very hard to heat up) or paper thin shell you still have to get rid of that energy.

You want to slow down as much as possible as high as possible (but I think SX are using some of the biggest Titanium forgings ever made for their first stage grid fins already). So your grid fins might have to be even bigger than on the booster stage, despite it being smaller and shorter. IOW recovery hardware does not necessarily scale with the length of the stage. 

High deceleration implies a lot of heat release as high speed air flow is brought to rest. Titanium starts to soften around 970c. That's hot by Aluminum standards but reentry airstream temps can hit 1800-2000c,

You could go with higher temperature materials but AFAIK all are much heavier. Titanium is about 50% heavier than Aluminum. Typical super alloys are 85% above Titanium. Beryllium is 2/3 the density of Aluminum and double the melting point of Al. It's also brittle, toxic and about 200x the cost.

And of course every unit of mass on the US comes off the payload 1:1, not about 13:1 as the booster enjoys. So despite the stage being smaller your drag generating system still ends up bigger and heavier than the booster, not what most people would expect.

Note that all of this is about the energy per unit mass you have to dissipate. This has to be done whatever shape the stage is, whatever attitude it comes in. It does not matter. The energy has to be dissipated, and it's about 4x higher per unit of stage weight than it was for the booster stag.

IOW The fact the US is smaller than the booster does not make the problem simpler.  It's the whole combination of factors, amount of energy, hard mass limits, complex control problems that probably explain why no one's made any serious go at US recovery so far.
« Last Edit: 09/02/2017 11:36 PM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline alang

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #105 on: 09/03/2017 12:26 AM »
What I don't know is what reducing orbital speed by an arbitrary percentage will do to the need for thermal protection. Perhaps there is some critical velocity that reentry needs to be below in order to avoid the need for more exotic materials.
You need a better understanding of what those terms mean for a start.

"orbital velocity" is the speed an object has to be moving around the Earth to balance the exact level of gravity at that altitude (the one you get when you plug the actual values for the earths radius and the object into the relevant formula).

But g rises as altitude falls, so once you slow down you will fall down.

So you can't be "reducing orbital speed by an arbitrary percentage" as you're not in orbit anymore, you're starting reentry.

Quote from: alang
If we assume that the upper stage separates later with F9 then I suppose the core stage also needs more fuel for its reentry burn to reduce the additional thermal stress. This all assumes that the side boosters carry the core stage so it doesn't burn so much fuel before they separate.
Or more TPS.

This problem is trickier than it looks. 

To come to land the stage has to lose basically all the energy it gained to get to orbit, which is a combination of kinetic and potential energy. Roughly speaking for S1 separation at 1/2 orbital velocity and 1/4 altitude (which keep the numbers simple) a unit of mass of the US has to lose 4x the energy a booster stage has to to land.

You can't decelerate completely using rocket thrust because you'll need as much propellant to slow down as you used to speed up. But even if you could bring it to a complete stop it will now fall 200Km straight down, picking up speed all the time.  Potential energy becomes kinetic energy and you still have to get rid of it.

This is on a "per mass" basis. IOW it doesn't matter if it's a small dense ball (very hard to heat up) or paper thin shell you still have to get rid of that energy.

You want to slow down as much as possible as high as possible (but I think SX are using some of the biggest Titanium forgings ever made for their first stage grid fins already). So your grid fins might have to be even bigger than on the booster stage, despite it being smaller and shorter. IOW recovery hardware does not necessarily scale with the length of the stage. 

High deceleration implies a lot of heat release as high speed air flow is brought to rest. Titanium starts to soften around 970c. That's hot by Aluminum standards but reentry airstream temps can hit 1800-2000c,

You could go with higher temperature materials but AFAIK all are much heavier. Titanium is about 50% heavier than Aluminum. Typical super alloys are 85% above Titanium. Beryllium is 2/3 the density of Aluminum and double the melting point of Al. It's also brittle, toxic and about 200x the cost.

And of course every unit of mass on the US comes off the payload 1:1, not about 13:1 as the booster enjoys. So despite the stage being smaller your drag generating system still ends up bigger and heavier than the booster, not what most people would expect.

Note that all of this is about the energy per unit mass you have to dissipate. This has to be done whatever shape the stage is, whatever attitude it comes in. It does not matter. The energy has to be dissipated, and it's about 4x higher per unit of stage weight than it was for the booster stag.

IOW The fact the US is smaller than the booster does not make the problem simpler.  It's the whole combination of factors, amount of energy, hard mass limits, complex control problems that probably explain why no one's made any serious go at US recovery so far.

Thanks for the detailed response. You're right I need to define my terms better.
However, regarding  "you'll need as much propellant to slow down as you used to speed up".
Not quite as the payload will be gone and so will a lot of the fuel.
I suspect SpaceX must be doing a lot of simulation work and the side boosters make everything more complicated.

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #106 on: 09/03/2017 08:47 AM »
Thanks for the detailed response. You're right I need to define my terms better.
However, regarding  "you'll need as much propellant to slow down as you used to speed up".
Not quite as the payload will be gone and so will a lot of the fuel.
As noted by nearly every beginner who's looked at this problem.  :(

The payload of a TSTO is about 2-3% of the whole rockets GTOW. A fully propulsive deceleration would need to cancel all the velocity the US has provided, and the velocity provided by the booster to the US as well.

Losing the payload is a very small part of the mass and velocity budget. You just spent 90%+ of your GTOW to get here, but now you have a propellant and engine that can cancel all of that, in a few % of the mass? Congratulations.

So why didn't you build the whole rocket to use that in the first place?

Quote from: alang
I suspect SpaceX must be doing a lot of simulation work and the side boosters make everything more complicated.
SX do a lot of simulation work and the FH side boosters will complicate recovery.
Simulation is important because any analogy should be treated with extreme suspicion. The results can be very counter intuitive.   :(

The numbers I chose kept the maths very simple. This threads only been talking about the F9, not the FH, but that splits the problem into 3 parts.

However the end part remains the same. Final stage is at orbital velocity. Recovery needs deceleration back to zero speed, zero altitude.

The only way that changes is if all payloads actively cooperate and supply the last bit of delta V to orbit. That guarantees the stage is actually sub-orbital, but it forces every payload to carry a rocket motor, which may be an unnecessary complication for a lot of them.

An obvious strategy for US recovery is "Just do what we did for booster recovery, it's basically the same problem."
Except it's not.  :( The generous exchange rate between recovery hardware and payload, choosing your staging velocity (to reduce heating) and lifetime of the stage are just a few things that you can't tweak if the numbers don't quite work out. Add in control surfaces that don't scale down as you might expect and may be much heavier anyway and what looks like a simple plan isn't.

This may be US recovery's biggest trap. It looks like a simpler (solved) problem, but it's not, especially if you're committed to doing it a certain way, as SX is. 

The premise of this thread, that you can use that information to feed into the design of ITS, seems like an extreme case of wishful thinking to me. I think this will be my last post.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline Xentry

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #107 on: 09/03/2017 03:26 PM »
Final stage is at orbital velocity. Recovery needs deceleration back to zero speed, zero altitude.
(...)
An obvious strategy for US recovery is "Just do what we did for booster recovery, it's basically the same problem."
Except it's not.  :(
(...)
The premise of this thread, that you can use that information to feed into the design of ITS, seems like an extreme case of wishful thinking to me. I think this will be my last post.

It's pretty clear that the Falcon 9 2nd stage recovery is essentially equivalent to the recovery of an orbital vehicle.
However, the cylindrical shape of the stage might be of some advantage relative to most other vehicles:
a) the area of the 2nd stage (assuming a side entry) would probably allow for a very small ballistic coefficient as compared to most Earth entry vehicles (perhaps even below 100kg/m2 - whereas those of Apollo and Dragon are higher than 300kg/m2 -, which is, coincidentally, more comparable to previous Mars entry vehicles than terrestrial ones...);
b) In that case, and because of the bottom-heavy configuration of propulsive stages, body flaps would be needed to force an entry with the required attitude for maximum drag (sideways). The body flaps could then also be used to ensure the production of lift, and possibly limited roll control authority (for cross-range);
c) A lifting entry with such a low beta means deceleration occurs at higher altitudes where density is much lower (at least 3-3.5x lower for the same deceleration). Heating would then be correspondingly lower (since its more than proportional to atmospheric density), and TPS needs would be lower too. I haven't done the simulations to see how low they could get, but am curious to see whether it is theoretically possible to fly trajectories using such a vehicle where maximum heat flux is brought down to values where titanium would be enough for protection, or where required TPS thickness would be so minimal as to essentially eliminate its' usual mass penalty;
d) with such a low ballistic coefficient, it might also make sense to descend sideways until a very low altitude, since terminal velocity would then be well below 50m/s (potentially about 35m/s). Body flaps could then be used to perform a kind of flare manoeuver, to restore a vertical attitude for the final, propulsive landing using a minimal amount of fuel.

Which brings me to a key point: this looks a lot like an EDL strategy which could be used by the BFS for a Mars landing, and it might even make sense to use a 2nd stage nearly as is to perform the first few tests whenever excess capacity is available (just add some body flaps to those flights in the same way the grid fins were added to the 1st stage, and add some TPS). This would not entail a complete redesign of the stage, and much less the construction of an ITSy...

Offline raketa

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #108 on: 09/06/2017 06:06 AM »
I also suspect the reuse of S2 is not primarily cost driven but for ITS development purposes.
What needs to be tested first are the hardest parts,
Raptor engine, landing cradle, refueling in orbit.
Heat shield also but there is probably less risk in it not working as envisioned.


Phase 0) Titanium grid fins on maybe a bit longer cylindrical S2

This configuration might not need a big heat shield to re-enter Earths atmosphere, just a lot of fuel, and much lower payload.

Phase 1) Landing in a landing cradle, might be possible with simple S2-gridfin, steering thrusters configuration but likely to fail several times (a Hail Mary attempt). It's better to use a S2 than an already re-usable S1 for these otherwise expensive landing cradle attempts. Grasshopper attempts will probably be done first.

If the cradle landing doesn't work, landing legs are needed, or super Draco engines for more a precise landing in the cradle

Phase 2) Raptor engine(s) for S2

Phase 3) Methalox S2 refueling in orbit

Refueling in orbit by another "tanker S2", might make heavier payloads and enough fuel for landing possible with this reusable S2 version.


Phase 4) Change the body to ITS Spaceship form

Landing Legs and Super Draco will make it work with more certainty but are non-essential for ITS development

Cradle landing
Great idea, the purpose will be not for testing first stage or ITSy, but save the weight of S2.
Grid fins will be able to steer it with feets precision and it will be good enough for landing S2.
Weight of modification could be lower under 2 tons.

Offline hkultala

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #109 on: 09/06/2017 06:13 AM »
I also suspect the reuse of S2 is not primarily cost driven but for ITS development purposes.
What needs to be tested first are the hardest parts,
Raptor engine, landing cradle, refueling in orbit.
Heat shield also but there is probably less risk in it not working as envisioned.


Phase 0) Titanium grid fins on maybe a bit longer cylindrical S2


Completely different direction than ITS re-entry.

Quote
This configuration might not need a big heat shield to re-enter Earths atmosphere, just a lot of fuel, and much lower payload.

"a lot of" is more fuel than it can carry, does not work.

Quote
Phase 1) Landing in a landing cradle, might be possible with simple S2-gridfin, steering thrusters configuration but likely to fail several times (a Hail Mary attempt). It's better to use a S2 than an already re-usable S1 for these otherwise expensive landing cradle attempts. Grasshopper attempts will probably be done first.

What engines do you plan to use for the landing itself? Merlin 1dvac is way too powerful with nozzle that is unstable in atmosphere, so it cannot be used for landing.

Offline alang

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #110 on: 09/09/2017 01:37 PM »
Parachutes didn't work for the recovery of the first stage. However, was that because they were used as the only option at first and not in conjunction with something else?
What about using steerable parachutes similar to but larger than those apparently being used for fairing reentry for the final part of second stage recovery onto land/bouncy castle? This reentry back onto land could be following one or more orbits and possible use of of some other form of stabilisation during and after reentry like cold gas thrusters and grid fins.
The point of this would be to recover enough of a second stage for it to be relevant for second stage quality improvements rather than immediate reuse and to test reentry technologies relevant to BFS. Even if it wasn't fully intact they still might learn something.

Online Eerie

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #111 on: 09/09/2017 04:29 PM »
Parachutes didn't work for the recovery of the first stage. However, was that because they were used as the only option at first and not in conjunction with something else?
What about using steerable parachutes similar to but larger than those apparently being used for fairing reentry for the final part of second stage recovery onto land/bouncy castle? This reentry back onto land could be following one or more orbits and possible use of of some other form of stabilisation during and after reentry like cold gas thrusters and grid fins.
The point of this would be to recover enough of a second stage for it to be relevant for second stage quality improvements rather than immediate reuse and to test reentry technologies relevant to BFS. Even if it wasn't fully intact they still might learn something.

S2 is re-entering from orbital speed. You can't use parachute for this. Otherwise, guess what, all the past and existing capsules would be using parachutes instead of heat shields.
« Last Edit: 09/09/2017 04:30 PM by Eerie »

Offline Welsh Dragon

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #112 on: 09/09/2017 06:22 PM »
Parachutes didn't work for the recovery of the first stage. However, was that because they were used as the only option at first and not in conjunction with something else?
What about using steerable parachutes similar to but larger than those apparently being used for fairing reentry for the final part of second stage recovery onto land/bouncy castle? This reentry back onto land could be following one or more orbits and possible use of of some other form of stabilisation during and after reentry like cold gas thrusters and grid fins.
The point of this would be to recover enough of a second stage for it to be relevant for second stage quality improvements rather than immediate reuse and to test reentry technologies relevant to BFS. Even if it wasn't fully intact they still might learn something.

S2 is re-entering from orbital speed. You can't use parachute for this. Otherwise, guess what, all the past and existing capsules would be using parachutes instead of heat shields.
Might want to read the whole post. Guess what, all past and existing capsules DO use parachutes for the final part of descent.

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #113 on: 09/09/2017 07:52 PM »
Parachutes didn't work for the recovery of the first stage. However, was that because they were used as the only option at first and not in conjunction with something else?
No. IIRC they were to be used with airbags, like the Kistler booster.
Quote from: alang
What about using steerable parachutes similar to but larger than those apparently being used for fairing reentry for the final part of second stage recovery onto land/bouncy castle? This reentry back onto land could be following one or more orbits
It won't be a couple of orbits. RTLS was the reason why the Shuttle was a double delta. Only an aircraft shape gave them the 1000s of Km of cross range needed to do this in 1 orbit. Boosters have very poor aerodynamics. It's very much simpler to just wait a day till it comes back over the launch site, but it needs more hardware to keep the stage alive for that length of time.
Quote from: alang
and possible use of of some other form of stabilisation during and after reentry like cold gas thrusters and grid fins.
So not an actual plan, more a set of random suggestions.  :(
You really should read some of the earlier posts on this thread.
Quote from: alang
The point of this would be to recover enough of a second stage for it to be relevant for second stage quality improvements rather than immediate reuse and to test reentry technologies relevant to BFS.
Except that they are not.  :(. AFAIK ITS is a stage shaped like a large lifting body, a very different beast from just a stage coming in at an angle. It's structurally very different. It'll use retro thrust for terminal velocity cancellation because parachutes don't work well on Mars because it's atmosphere is 1/160 that of Earth.
Quote from: alang
Even if it wasn't fully intact they still might learn something.
Highly unlikely to make the cost of the modification worthwhile.

You'd really be well advised to read the last 2-3 pages of this thread before commenting further.  :(
« Last Edit: 09/09/2017 07:55 PM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline meekGee

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #114 on: 09/10/2017 03:52 AM »
Parachutes didn't work for the recovery of the first stage. However, was that because they were used as the only option at first and not in conjunction with something else?
No. IIRC they were to be used with airbags, like the Kistler booster.
Quote from: alang
What about using steerable parachutes similar to but larger than those apparently being used for fairing reentry for the final part of second stage recovery onto land/bouncy castle? This reentry back onto land could be following one or more orbits
It won't be a couple of orbits. RTLS was the reason why the Shuttle was a double delta. Only an aircraft shape gave them the 1000s of Km of cross range needed to do this in 1 orbit. Boosters have very poor aerodynamics. It's very much simpler to just wait a day till it comes back over the launch site, but it needs more hardware to keep the stage alive for that length of time.
Quote from: alang
and possible use of of some other form of stabilisation during and after reentry like cold gas thrusters and grid fins.
So not an actual plan, more a set of random suggestions.  :(
You really should read some of the earlier posts on this thread.
Quote from: alang
The point of this would be to recover enough of a second stage for it to be relevant for second stage quality improvements rather than immediate reuse and to test reentry technologies relevant to BFS.
Except that they are not.  :(. AFAIK ITS is a stage shaped like a large lifting body, a very different beast from just a stage coming in at an angle. It's structurally very different. It'll use retro thrust for terminal velocity cancellation because parachutes don't work well on Mars because it's atmosphere is 1/160 that of Earth.
Quote from: alang
Even if it wasn't fully intact they still might learn something.
Highly unlikely to make the cost of the modification worthwhile.

You'd really be well advised to read the last 2-3 pages of this thread before commenting further.  :(

What you mean is that he's not agreeing with your posts from the last 2-3 pages.

You keep thinking of S2 recovery in terms of S1 recovery, and then keep showing how impossible that is.

But S2 recovery has nothing to do with S1 recovery.

Look at the fundamentals:

S1
Speed:  1.5 km/s
Mass: 22 tons
Aspect Ratio: 11.5:1
Landing location:  Constrained

S2
Speed:  Orbital
Mass:  4 Ton
Aspect Ratio:  2.5:1  (or 3.5:1 w/nozzle)
Landing location:  Free

Dragon 2
Speed:   Orbital
Mass:  6 Ton
Aspect Ratio:  1:1
Landing location:  Free

So a reasonable plan would be:
1) Add heat shield and a stabilizing drag device for passive reentry
2) Possibly add grid fins for trajectory adjustment
3) Add parachute
4) detach nozzle before reentry (mechanical connection only)
5) Mid-air recovery just off-shore

And allow up to 4 tons for the combined hardware. (Heat Shield, drag devices, parachute)

Seems a lot more reasonable than carrying propellant to slow down from orbital speed which is probably 30 tons (based on 100 tons of propellant in the up direction and 1/3 empty mass), plus landing engines, plus legs
« Last Edit: 09/10/2017 08:51 AM by meekGee »
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Offline wannamoonbase

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #115 on: 09/10/2017 03:09 PM »
MeekGee,

I agree with mid air recovery. 

Most mass efficient and more freedom in recovery zones. 

I think there will also need to be some type of hardware added that manages the US for sometime till it can be over the right recovery area.
Excited to be finally into the first Falcon Heavy flow, we are getting so close!

Online douglas100

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #116 on: 09/10/2017 04:30 PM »
Also agree that it's a sensible approach. Essentially the stage only needs to stay alive 24 hours or so. Batteries should handle that. LOX boil off shouldn't be a problem for a stage in LEO. De-orbit can be handled by a set of Dracos if necessary. Also useful for control during entry. It might not even be necessary to eject the engine nozzle. After the helicopter has caught the stage it could be lowered horizontally into a cradle on the ground so that the nozzle isn't impacted.
Douglas Clark

Offline MikeAtkinson

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #117 on: 09/10/2017 06:02 PM »
Also agree that it's a sensible approach. Essentially the stage only needs to stay alive 24 hours or so.

~12 hours I believe, until landing site rotates under the far part of the orbit from the launch site.

Whether landing can be made at 12 or 24 hours, depends to some extent on the orbit payload inserted into, subsequent S2 orbit changes and S2 crossrange, and on flexibility of landing position. Midair recovery aids a flexible landing position. Recovery away from the launch site enables much shorter flights, potentially only half an orbit or so with recovery near western Australia.

Landing on a specific pad at the launch site could take many days if crossrange is low and there S2 orbit changes are limited.

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #118 on: 09/11/2017 07:32 AM »

S1
Speed:  1.5 km/s
Mass: 22 tons
Aspect Ratio: 11.5:1
Landing location:  Constrained

S2
Speed:  Orbital
Mass:  4 Ton
Aspect Ratio:  2.5:1  (or 3.5:1 w/nozzle)
Landing location:  Free

Dragon 2
Speed:   Orbital
Mass:  6 Ton
Aspect Ratio:  1:1
Landing location:  Free
I deeply dislike reasoning by analogy. Once you abandon the idea that US recovery is not going to work like booster recovery you are faced with coming up with a different scheme.

Let me see if we can agree on 2 things.
1)Roughly speaking a US had to lose 4x the energy per unit mass that a booster has to lose to land.
2)The "exchange rate" is much less forgiving for US than booster.

Can we do that? Let's see what  you've come up with.
Quote from: meekGee
So a reasonable plan would be:
1) Add heat shield and a stabilizing drag device for passive reentry
Except that's not a plan, is it? That's a wish list for something you hope can be done.
An actual plan would suggest what that would be and how it would be attached to the stage.

Both stages are alike in that they will be bottom heavy so they will tend to flip heavy end on to the flow. S1 recovery leaves it there but the US  has 4x as much energy to lose (per unit mass) as the booster.

BTW "We're not doing it like booster recovery" doesn't get you out of trouble. If you decided to land the booster the same way as the US any place you put PICAX on the US could be 1/4 the thickness on the booster, but you've got the exchange rate working for you there, so it's not likely to be an issue.

Quote from: meekGee
2) Possibly add grid fins for trajectory adjustment
Which may burn up if they are Titanium and are not going to be based on stage size. Titanium softens at 970c (of course actual peak operating temp will depend on applied loads, 970c is minimum forging temp when being whacked with a hammer) . Anything above that and you're looking to upgrade to an Iron based high temp steel or a Nickel based super alloy (at least 73% heavier, even if the same size), or (worse case) go to an actual refractory element like Molybdenum or Niobium (Good news. Nb is only 83% denser than Ti) .
Quote from: meekGee
3) Add parachute

Again, where? Top end, bottom end or on the side? How much reinforcement will this stage need to stop being torn apart by the shock loads?
Quote from: meekGee
4) detach nozzle before reentry (mechanical connection only)

5) Mid-air recovery just off-shore
How does this reference back to improving the design of ITS, which is the supposed driving idea for this thread?
Quote from: meekGee
And allow up to 4 tons for the combined hardware. (Heat Shield, drag devices, parachute)

Seems a lot more reasonable than carrying propellant to slow down from orbital speed which is probably 30 tons (based on 100 tons of propellant in the up direction and 1/3 empty mass), plus landing engines, plus legs
Realistically no stage can do that anyway.  :( . That pesky "every m/s you put on with a Kg of propellant you have to take off with another Kg of propellant" rule.
 Shuttle OMS provided a delta v of maybe 100m/s to get the ball rolling.
Everything else was air friction and wing generated lift. In the case of F9 S1 it's mostly air friction. 

You're right that works for booster recovery won't work for US recovery, which I'd say SX have known since at least 2014.

Perhaps you could come up with an actual plan?
« Last Edit: 09/11/2017 07:42 AM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline meekGee

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #119 on: 09/11/2017 07:43 AM »

I deeply dislike reasoning by analogy. Once you abandon the idea that US recovery is not going to work like booster recovery you are faced with coming up with a different scheme.

Let me see if we can agree on 2 things.
1)Roughly speaking a US had to lose 4x the energy per unit mass that a booster has to lose to land.
2)The "exchange rate" is much less forgiving for US than booster.

Can we do that? Let's see what  you've come up with.

....
....
....

You're right that works for booster recovery won't work for US recovery, which I'd say SX have known since at least 2014.

Perhaps you could come up with an actual plan?

See, I never "abandoned" the idea that US recovery would work like booster recovery, since I've never had it in the first place.  You're the only one who had it.

Instead, I look at the closer analog I can find, for a stubby orbital object with a mass of approximately 5 tons, and I see Dragon, or capsules in general.  I don't even consider the booster, since that's an entirely different problem, from any angle I look at it.

You say that "You're right that works for booster recovery won't work for US recovery, which I'd say SX have known since at least 2014."

Actually, I think they've known that since forever, since the thought of doing S1-like recovery for S2 never crossed their minds. 

The reusability video showed ballistic re-entry, but used a landing engine/legs instead of parachutes.  THAT may have changed, but that's nowhere near doing boost-back and re-entry burns.

As for an "actual plan" - what exactly are you expecting?  My engineering staff is currently pre-occupied with several other projects.
« Last Edit: 09/11/2017 08:27 AM by meekGee »
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Offline DreamyPickle

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #120 on: 09/11/2017 08:40 AM »
Quote
Both stages are alike in that they will be bottom heavy so they will tend to flip heavy end on to the flow.

Yes, this is a real problem. Significant effort went in properly designing the center-of-mass for the shuttle, X-33 and similar vehicles. The BFS will also have the same problem as F9 S2: It will be extremely bottom heavy.

However if you look at http://BFS drawings you see two spherical tanks at the top of each propellant tank. If the landing propellant is moved in there then the center of mass will move forward quite a bit. It also has the advantage that by having a smaller surface area it heats up slower.



I don't know how to adapt this for F9.
« Last Edit: 09/11/2017 08:41 AM by DreamyPickle »

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #121 on: 09/11/2017 11:23 AM »
I don't know how to adapt this for F9.
That's really the core of the problem with this thread. :(  Even if you can do all this to an F9 US how does that feed into ITS, or vice versa.

Small factoid.

The concept of "sump tanks" was used in the Kistler RLV design but AFAIK dates back to bowl shaped depressions in the tank ends of the Agena stage to encourage enough propellant to be retained by surface tension forces to smooth engine restart. NASA has also done work to improve fluid capture in tank feed lines by (effectively) tapering them, but historically this has been expensive to do. IIRC sounding rocket tests showed a 50% improvement in collecting fluid, not vapor.

Agena was a very impressive design. Part stage, part satellite bus and probably the only US hypergolic stage design with common bulkhead tanks, yet AFAIK then never failed.   :o
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline rsdavis9

1)Roughly speaking a US had to lose 4x the energy per unit mass that a booster has to lose to land.


isn't the energy per unit mass 22 times?
(7/1.5)^2
bob

Offline wannamoonbase

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #123 on: 09/11/2017 01:18 PM »
1)Roughly speaking a US had to lose 4x the energy per unit mass that a booster has to lose to land.


isn't the energy per unit mass 22 times?
(7/1.5)^2


Kinetic Energy =  (4/1.5)^2 = 21.8 times

Heat Energy Power= (7/1.5)^3 = 101.6 times
« Last Edit: 09/11/2017 02:09 PM by wannamoonbase »
Excited to be finally into the first Falcon Heavy flow, we are getting so close!

Online Robotbeat

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #124 on: 09/11/2017 02:03 PM »
Heat power, not energy.
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Offline rsdavis9

Quote
Kinetic Energy =  (4/1.5)^2 = 21.8 times

Heat Energy = (7/1.5)^3 = 101.6 times
Sorry if I am being a dummy here but energy = heat energy.
Where did the 4 come from in the 4/1.5 shouldn't it be 7/1.5
Kinetic energy will turn into heat one for one.

EDIT:
Sorry I missed the correction of heat energy to heat power.
So why is Heat Power a cubed relation? Couldn't find any reference...
Probably something to do with hypersonic compression of a gas?
« Last Edit: 09/11/2017 03:03 PM by rsdavis9 »
bob

Offline meekGee

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #126 on: 09/11/2017 03:03 PM »
Quote
Both stages are alike in that they will be bottom heavy so they will tend to flip heavy end on to the flow.

Yes, this is a real problem. Significant effort went in properly designing the center-of-mass for the shuttle, X-33 and similar vehicles. The BFS will also have the same problem as F9 S2: It will be extremely bottom heavy.

However if you look at http://BFS drawings you see two spherical tanks at the top of each propellant tank. If the landing propellant is moved in there then the center of mass will move forward quite a bit. It also has the advantage that by having a smaller surface area it heats up slower.



I don't know how to adapt this for F9.
Of course it is, but yes, it is solvable.

Very little fuel is left over after the reentry burn - only enough for the landing burn which in our case is non existent, so any contribution of fuel to the problems is non existent as well.

Heat shield and parachute  both  near the front end, will help with the c.g.  So will dropping the nozzle.

Drag device will help with the c.d.

Maybe the active control device is a sled, located near the heat shield.

Point is, if we're allowing up to 100% mass penalty, c.g. location can't be a show stopper.
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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #127 on: 09/11/2017 04:31 PM »
Quote
Kinetic Energy =  (4/1.5)^2 = 21.8 times

Heat Energy = (7/1.5)^3 = 101.6 times
Sorry if I am being a dummy here but energy = heat energy.
Where did the 4 come from in the 4/1.5 shouldn't it be 7/1.5
Kinetic energy will turn into heat one for one.

EDIT:
Sorry I missed the correction of heat energy to heat power.
So why is Heat Power a cubed relation? Couldn't find any reference...
Probably something to do with hypersonic compression of a gas?
Think of the mass of air that the cross section displaces. That displaced mass flow per unit time is proportional to velocity, and the energy dumped per unit mass is proportional to velocity squared (i.e. kinetic energy). Multiply those together and you get cubed.

Nothing at all complicated about this.
« Last Edit: 09/11/2017 04:34 PM by Robotbeat »
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Offline rsdavis9

Quote
Kinetic Energy =  (4/1.5)^2 = 21.8 times

Heat Energy = (7/1.5)^3 = 101.6 times
Sorry if I am being a dummy here but energy = heat energy.
Where did the 4 come from in the 4/1.5 shouldn't it be 7/1.5
Kinetic energy will turn into heat one for one.

EDIT:
Sorry I missed the correction of heat energy to heat power.
So why is Heat Power a cubed relation? Couldn't find any reference...
Probably something to do with hypersonic compression of a gas?
Think of the mass of air that the cross section displaces. That displaced mass flow per unit time is proportional to velocity, and the energy dumped per unit mass is proportional to velocity squared (i.e. kinetic energy). Multiply those together and you get cubed.

Nothing at all complicated about this.

As einstein said if you can't explain it to your grandmother than you really don't understand it.
Thats why I like the simple first order "feel" explanations.

Thank you
bob

Offline envy887

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #129 on: 09/11/2017 08:14 PM »
Also agree that it's a sensible approach. Essentially the stage only needs to stay alive 24 hours or so. Batteries should handle that. LOX boil off shouldn't be a problem for a stage in LEO. De-orbit can be handled by a set of Dracos if necessary. Also useful for control during entry. It might not even be necessary to eject the engine nozzle. After the helicopter has caught the stage it could be lowered horizontally into a cradle on the ground so that the nozzle isn't impacted.

Doesn't even need to stay alive for 24 hours. For a typical GTO or ISS launch, it will pass over the Pacific within 1000 miles of California in 2 to 10 hours. Almost everything out of Vandy can once around 1200 to 1800 miles out in the Pacific within 2 hours. Just need a ship out of LA that can support a largish helicopter. Grab the stage under chutes, drop it in a large airbag on the deck, and bring everything home.

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #130 on: 09/11/2017 10:29 PM »
1)Roughly speaking a US had to lose 4x the energy per unit mass that a booster has to lose to land.


isn't the energy per unit mass 22 times?
(7/1.5)^2
Note that word "roughly."  :(

To keep the maths really simple on the differences between S2 and S1 I took stage separation at 1/2 orbital velocity and altitude at 1/4 altitude. Hence "roughly" 4x difference. it kept the numbers simple enough to not need a calculator once you included both KE and PE.

But in fact the real story is even worse.  :( At 200Km it's nearer 7794m/s roughly 27:1 on KE.

I used the crude 4x because Shoywell said the current generation of F9 boosters were good for about 3 launches, so if the environment was 4x worse then doing what you did on the booster in terms of TPS thickness would not survive 1 use, and historically TSTO ELV's have split the delta V roughly in 2 equal pieces.

In fact (as you've noted ) the orbital situation is so much worse than the boost stage because booster MECO is well below 1/2 orbital speed.

IOW SX will need a TPS that's 27x better than the stuff on the booster per unit mass.
Most people don't expect that big a hit. Any area that's got say 1lb of TPS on the booster (and is equally exposed on the US) will need 27lb of TPS to protect it.

I'm pretty sure the science doesn't have that much stretch left in it.

[EDIT. TPS "design" is a bit of a misnomer. The composition, working and heat treatment of the materials involved can have such a huge effect on the final properties that it's more like baking. Just as flour, fat, salt and water can give you bread, pizza, flaky pastry or pie casing. Same ingredients, very different outcomes.]
« Last Edit: 09/12/2017 06:40 PM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #131 on: 09/11/2017 10:45 PM »
See, I never "abandoned" the idea that US recovery would work like booster recovery, since I've never had it in the first place.  You're the only one who had it.
And you don't know much about how engineers actually do their job either.

Generally if they find something that works, especially after a long hard search with dead ends, they will try to apply it as widely as possible to situations, unless it is very inefficient. I'd be very surprised if SX didn't run some simulations to double check if the S1 recovery plan could work for US use.
Quote from: meekGee
Instead, I look at the closer analog I can find, for a stubby orbital object with a mass of approximately 5 tons, and I see Dragon, or capsules in general.  I don't even consider the booster, since that's an entirely different problem, from any angle I look at it.
Which is short, very dense and can put all its structure in the shadow of the heat shield, which is impossible for a stage.
Quote from: meekGee
You say that "You're right that works for booster recovery won't work for US recovery, which I'd say SX have known since at least 2014."

Actually, I think they've known that since forever, since the thought of doing S1-like recovery for S2 never crossed their minds. 
You're entitled to your opinion. Perhaps you should take another look at the video SX released in 2011?
Quote from: meekGee
The reusability video showed ballistic re-entry, but used a landing engine/legs instead of parachutes.  THAT may have changed, but that's nowhere near doing boost-back and re-entry burns.
If you're already in orbit you don't have to do "boost back" as you're "coming round again" on the launch site. What you do need to do is decelerate enough to start the reentry process. Shuttle did it exactly the way the S1 does it, retro burn (using OMS) but then flipped back over for nose first entry as it would have been difficult to land rear end first.
Quote from: meekGee
As for an "actual plan" - what exactly are you expecting?  My engineering staff is currently pre-occupied with several other projects.
Something more than a wish list?

Wishing is for wizards who can cast spells.  :( I haven't found magic to work in this universe, so I tend to rely on engineering. 
« Last Edit: 09/12/2017 07:52 AM by john smith 19 »
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline speedevil

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #132 on: 09/14/2017 12:53 PM »
https://twitter.com/elonmusk/status/908254079092002816
Quote

Long road to reusabity of Falcon 9 primary boost stage…When upper stage & fairing also reusable, costs will drop by a factor >100.

If F9 is going to be reusable with very high cycle counts, it implies that ITSy might be somewhat delayed?
Or could this in fact act to advance things, if revenue from the constellation comes online sooner than expected.



« Last Edit: 09/14/2017 01:00 PM by speedevil »

Offline AncientU

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #133 on: 09/14/2017 02:57 PM »
https://twitter.com/elonmusk/status/908254079092002816
Quote

Long road to reusabity of Falcon 9 primary boost stage…When upper stage & fairing also reusable, costs will drop by a factor >100.

If F9 is going to be reusable with very high cycle counts, it implies that ITSy might be somewhat delayed?
Or could this in fact act to advance things, if revenue from the constellation comes online sooner than expected.

F9 fairing and second stage reuse are already in flight test (or late development); means that ITSy may be somewhat advanced.  Think of all this kaboomy goodness as prototyping for ITSy.
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Offline su27k

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #134 on: 09/15/2017 03:59 AM »
https://twitter.com/elonmusk/status/908254079092002816
Quote

Long road to reusabity of Falcon 9 primary boost stage…When upper stage & fairing also reusable, costs will drop by a factor >100.

If F9 is going to be reusable with very high cycle counts, it implies that ITSy might be somewhat delayed?
Or could this in fact act to advance things, if revenue from the constellation comes online sooner than expected.

If they want to start deploying constellation in 2019, they'll have to rely on F9, so a reusable S2 is a good investment, I don't think ITSy is part of the equation.

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #135 on: 09/15/2017 04:33 AM »
They want Falcon 9 highly reusable so they can shut down the Falcon production line and switch it over to BFR.
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Offline woods170

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #136 on: 09/15/2017 05:54 AM »
They want Falcon 9 highly reusable so they can shut down the Falcon production line and switch it over to BFR.
No. the Falcon production line will stay open, albeit down-sized. BFR and Falcon 9 will co-exist for quite some time with Falcon 9 being the working horse (needing fresh stages once in a while to replace warn-out reused ones) whilst SpaceX is working the bugs out of BFR.

Offline Lars-J

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #137 on: 09/15/2017 06:39 AM »
They want Falcon 9 highly reusable so they can shut down the Falcon production line and switch it over to BFR.
No. the Falcon production line will stay open, albeit down-sized. BFR and Falcon 9 will co-exist for quite some time with Falcon 9 being the working horse (needing fresh stages once in a while to replace warn-out reused ones) whilst SpaceX is working the bugs out of BFR.

Don't forget about the Falcon upper stages. They are build on the same production line as the first stages, so they will stay busy. (And reuse of falcon upper stages is far off, even if they will experiment)

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #138 on: 09/15/2017 07:03 AM »
No. the Falcon production line will stay open, albeit down-sized. BFR and Falcon 9 will co-exist for quite some time with Falcon 9 being the working horse (needing fresh stages once in a while to replace warn-out reused ones) whilst SpaceX is working the bugs out of BFR.
Agreed. Lots of people have noted that SX does not retain production of anything that's served its purpose but IRL F9 is the design that's launching payloads right now.  FH is 2/3s more or less stock F9s and that's not even flown yet.  I doubt it will reach RL10/Centaur levels of longevity (is there any other US design from that era still flying?) but it seems safe to say it will be around for at least a decade.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline livingjw

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #139 on: 09/15/2017 11:39 AM »
I suspect that a reusable upper stage for F9 would have an internal layout where the cargo is placed between the propellant tanks. This would place the payload near the cg. Something like the X-37B but with a wingless lifting body similar to ITS.

John

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #140 on: 09/15/2017 11:43 AM »
They want Falcon 9 highly reusable so they can shut down the Falcon production line and switch it over to BFR.
No. the Falcon production line will stay open, albeit down-sized. BFR and Falcon 9 will co-exist for quite some time with Falcon 9 being the working horse (needing fresh stages once in a while to replace warn-out reused ones) whilst SpaceX is working the bugs out of BFR.
Theyll coexist for a while, which is why they're trying to make Falcon 9 highly reusable, even the upper.

Build a stockpile of boosters and stages to last until retirement then shut down production.
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Offline AncientU

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #141 on: 09/15/2017 12:02 PM »
They want Falcon 9 highly reusable so they can shut down the Falcon production line and switch it over to BFR.
No. the Falcon production line will stay open, albeit down-sized. BFR and Falcon 9 will co-exist for quite some time with Falcon 9 being the working horse (needing fresh stages once in a while to replace warn-out reused ones) whilst SpaceX is working the bugs out of BFR.
Theyll coexist for a while, which is why they're trying to make Falcon 9 highly reusable, even the upper.

Build a stockpile of boosters and stages to last until retirement then shut down production.

Stockpile and shut down production seem like pure speculation... has anything ever been mentioned about stockpiling or an end to Falcon production? 

Note: The comment about Raptor-powered launcher making all other rockets obsolete is fairly generic and speculative -- like the thousands of massive satellites per year that was mentioned in same interview.
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Offline alang

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #142 on: 09/17/2017 12:56 PM »
I suspect that a reusable upper stage for F9 would have an internal layout where the cargo is placed between the propellant tanks. This would place the payload near the cg. Something like the X-37B but with a wingless lifting body similar to ITS.

John

It's a shame that none of the grown ups responded to this. Maybe it's too 'out there' and would require a lot of work. What are the problems with the plumbing for the propellant being separated like that? Has anyone ever done it?

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #143 on: 09/17/2017 01:24 PM »
Except cargo needs a big space, like the current fairings. No way there'd be enough room between the tanks, and if there was, you'd make the upper stage way too heavy due to long, heavy structural members between the tanks.
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Offline rsdavis9

Well if the cargo bay is like a big tank it shouldn't need any more structure. Just a lid on one side. Maybe the lid could be like 1/3 of the circumference? Thereby retaining most of the strength of the cylinder. Maybe the lid has latches that distribute the load? The cargo bay shouldn't be heavier than the current fairing.
bob

Offline Patchouli

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #145 on: 09/17/2017 06:00 PM »
Well if the cargo bay is like a big tank it shouldn't need any more structure. Just a lid on one side. Maybe the lid could be like 1/3 of the circumference? Thereby retaining most of the strength of the cylinder. Maybe the lid has latches that distribute the load? The cargo bay shouldn't be heavier than the current fairing.

If it has support the mass of a tank in front it would need be heavier as it has to be stronger.
Though this may not be a deal killer for many missions.
A possible evolution might be to make the F9 upper stage into something like the Kitsler K1 OTV.

One bullet Musk might have to bite is accepting use of a parachute and airbag landing system as powered landings are probably going to be out or the question though it could be redesigned as a stub winged space plane similar to the North American DC-3 or a system similar the ESA Adeline.
« Last Edit: 09/17/2017 06:08 PM by Patchouli »

Offline wannamoonbase

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #146 on: 09/17/2017 06:13 PM »
One bullet Musk might have to bite is accepting use of a parachute and airbag landing system as powered landings are probably going to be out or the question though it could be redesigned as a stub winged space plane similar to the North American DC-3 or a system similar the ESA Adeline.

Parachute and air capture with a heavy lift helicopter so that it can be set down gently would have the least weight impact.

Not sure how practical that is with weight of the US and payload and range of a chopper.  Maybe test over Edwards or off the east or west coast and set down in a port.
Excited to be finally into the first Falcon Heavy flow, we are getting so close!

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #147 on: 09/17/2017 06:24 PM »
Except cargo needs a big space, like the current fairings. No way there'd be enough room between the tanks, and if there was, you'd make the upper stage way too heavy due to long, heavy structural members between the tanks.
I'd say one of the grown ups just has.
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline cambrianera

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #148 on: 09/17/2017 06:38 PM »
Well if the cargo bay is like a big tank it shouldn't need any more structure. Just a lid on one side. Maybe the lid could be like 1/3 of the circumference? Thereby retaining most of the strength of the cylinder. Maybe the lid has latches that distribute the load? The cargo bay shouldn't be heavier than the current fairing.

The axial load carrying element in a tank is...
the pressurizing gas inside.
Oh to be young again. . .

Offline john smith 19

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #149 on: 09/17/2017 07:48 PM »
The axial load carrying element in a tank is...
the pressurizing gas inside.
No, no, it's alright. The lid (despite being very large and subject to temperature variations of several 100c) will be airtight.

And once the payload is ejected it'll seal perfectly and the payload volume be occupied by gas at several atm (IIRC the Atlas pressure was about 80psi on the ground, so it will be lower on orbit but have to rise to give the stiffness).

Oh yes, this will be so easy people will wonder why it's never been done before.  ;)
"Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry Spencer 1/28/11  Averse to bold? You must be in marketing."It's all in the sequencing" K. Mattingly.  STS-Keeping most of the stakeholders happy most of the time.

Offline meekGee

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #150 on: 09/17/2017 08:10 PM »
The axial load carrying element in a tank is...
the pressurizing gas inside.
No, no, it's alright. The lid (despite being very large and subject to temperature variations of several 100c) will be airtight.

And once the payload is ejected it'll seal perfectly and the payload volume be occupied by gas at several atm (IIRC the Atlas pressure was about 80psi on the ground, so it will be lower on orbit but have to rise to give the stiffness).

Oh yes, this will be so easy people will wonder why it's never been done before.  ;)

The extra strength is mostly needed on ascent.  So the seal only has to be effective "as launched".

Temperature difference of 100s of degrees - only if you don't try to mitigate them, and there are plenty of ways to do so.

Clearly a payload bay door is not easy, but it doesn't require magic.

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Offline AncientU

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #151 on: 09/17/2017 08:56 PM »
Well if the cargo bay is like a big tank it shouldn't need any more structure. Just a lid on one side. Maybe the lid could be like 1/3 of the circumference? Thereby retaining most of the strength of the cylinder. Maybe the lid has latches that distribute the load? The cargo bay shouldn't be heavier than the current fairing.

The axial load carrying element in a tank is...
the pressurizing gas inside.

Don't believe this is true.  Axial load can only be carried by a rigid member, not a gas.  The skin, when pressurized, is maintained as a right circular cylinder and buckling failure is prevented by pressurization.
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Offline gin455res

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #152 on: 09/17/2017 09:33 PM »
Is there any sense in having the f9-heavy central core stage and the upper-stage the same length as below (modified from image on Wikipedia):


The core stage would be a Falcon-5 with legs and the upper-stage a falcon-3 (1 standard merlin + 2 twin-chamber over-expanded merlins).


This might allow attempting cradle landings of the upper-stage.


Ps. I think it is worth remembering that at the initiation of re-entry the upper-stage is weightless and the angle of entry is very small.  (I imagine the above upper-stage entering base first)

Offline livingjw

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #153 on: 09/18/2017 12:54 AM »
Except cargo needs a big space, like the current fairings. No way there'd be enough room between the tanks, and if there was, you'd make the upper stage way too heavy due to long, heavy structural members between the tanks.

Except, by the time we add the weight needed to recover S2, you won't have near the payload capacity of the expendable S2 and hence, will not need such large volumes.  In fact you may not have much more payload capacity than the Dragon. We need to find a way to:

- Move the cg forward
- Move the AC back (without adding a lot of surface area)
- Keep the heat shield sealed until after reentry.

To me, this says cargo bay doors with a fixed nose cone.

John

Offline KelvinZero

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #154 on: 09/18/2017 02:15 AM »
Hi, in that ITS upper stage schematic, is oxygen on top or below? Just wondering since oxygen is so much more significant mass wise than the methane I think. (an online example had 8g methane burning with 32g oxygen)

Online Semmel

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #155 on: 09/18/2017 05:02 AM »
LOX was above in the first stage, can't remember for the second. The feed pipe for the engines was used as pressurised landing-fuel tank so no inside sphere required.

Offline hkultala

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #156 on: 09/18/2017 05:35 AM »
Is there any sense in having the f9-heavy central core stage and the upper-stage the same length as below (modified from image on Wikipedia):


The core stage would be a Falcon-5 with legs and the upper-stage a falcon-3 (1 standard merlin + 2 twin-chamber over-expanded merlins).

So many different new stages and even a new engine type for minimal(or negative) benefit. Costs would skyrocket, absolutely no sense.

And the much longer second stage would just be more DIFFICULT to survive the the re-entry than current second stage.

What is the point of the hyphothetical "twin-chamber over-expanded engine"? That is the part that adds the "absolutely" part into my judgement.

Quote

This might allow attempting cradle landings of the upper-stage.

No, it would not make it any easier, would only make it harder.
« Last Edit: 09/18/2017 06:24 AM by hkultala »

Offline su27k

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #157 on: 09/18/2017 05:41 AM »
I suspect that a reusable upper stage for F9 would have an internal layout where the cargo is placed between the propellant tanks. This would place the payload near the cg. Something like the X-37B but with a wingless lifting body similar to ITS.

John

It's a shame that none of the grown ups responded to this. Maybe it's too 'out there' and would require a lot of work. What are the problems with the plumbing for the propellant being separated like that? Has anyone ever done it?

I'm confused, why are we even thinking about payload bay on Falcon 9? Seems a complete non-starter, for one SpaceX is working on fairing reuse, this would negate the use of a payload bay. For another, a payload bay would be too small for many payloads, which means you'll need a 2nd stage without payload bay to handle those, so basically two types of second stages, two production lines, seems way too expensive and time consuming when SpaceX is supposed to be moving on to ITS next.

Online guckyfan

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #158 on: 09/18/2017 05:52 AM »
I'm confused, why are we even thinking about payload bay on Falcon 9? Seems a complete non-starter, for one SpaceX is working on fairing reuse, this would negate the use of a payload bay. For another, a payload bay would be too small for many payloads, which means you'll need a 2nd stage without payload bay to handle those, so basically two types of second stages, two production lines, seems way too expensive and time consuming when SpaceX is supposed to be moving on to ITS next.

I agree. No payload bay on Falcon upper stages. But there will be a need for 2 types of second stage. There will be a number of flights where the second stage can not be reused. Probably anything beyond GTO and everything that circularizes at higher than LEO. So the old type will need to remain available unless there all reuse hardware is only add on and the basic design remains the same.

Offline gin455res

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #159 on: 09/18/2017 07:11 AM »
Is there any sense in having the f9-heavy central core stage and the upper-stage the same length as below (modified from image on Wikipedia):


The core stage would be a Falcon-5 with legs and the upper-stage a falcon-3 (1 standard merlin + 2 twin-chamber over-expanded merlins).

So many different new stages and even a ne engine type for minimal benefit. Costs would skyrocket, absolutely no sense.

What is the point of the hyphothetical "twin-chamber over-expanded engine"? That is the part that adds the "absolutely" part into my judgement.

Quote

This might allow attempting cradle landings of the upper-stage.

No, it would not make it any easier, would only make it harder.


Yes there are new stages. The Falcon 5 is a shortened Falcon-9 with 4 engines removed. The upper-stage is then based on the same tankage.


The upper-stage is much thinner/finer than the current one and might be more stable re-entering.  It would re-enter base first, be quite fluffy and maybe better shaped for hypersonic flight than a more stumpy stage.


It would have more rotational inertia. It is easier to balance a broom on your hand than a pencil. The US has a central sea-level engine that can be used for landing/cradling. 


The twin chamber engines have better isp than the regular merlins for acceptable upper-stage economy. 2 x fat-nozzled twin-chamber engines  + 1 regular engine allow 3 engines to pack better under a stage. The four fat nozzles arranged in an exploded square surrounding the smaller sea-level engine.


Do we know how robust the current vacuum merlin nozzle would be through re-entry?


The combined stack will be shorter (FH is quite thin/fine.)


Would you prefer two stacked Falcon-3-s for more commonality, or a Falcon-5 with a stretched conventional upper-stage


(new tangent: would a four chambered vacuum merlin - like the rd-0124 - be more robust through re-entry?)

[/quote]

Offline cambrianera

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #160 on: 09/18/2017 07:58 AM »
Well if the cargo bay is like a big tank it shouldn't need any more structure. Just a lid on one side. Maybe the lid could be like 1/3 of the circumference? Thereby retaining most of the strength of the cylinder. Maybe the lid has latches that distribute the load? The cargo bay shouldn't be heavier than the current fairing.

The axial load carrying element in a tank is...
the pressurizing gas inside.

Don't believe this is true.  Axial load can only be carried by a rigid member, not a gas.  The skin, when pressurized, is maintained as a right circular cylinder and buckling failure is prevented by pressurization.

Something as simple as a pneumatic cylinder shows gas is used as an axial load carrying element so often.
Or do you prefer tyres?
Oh to be young again. . .

Offline livingjw

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #161 on: 09/18/2017 11:44 AM »
I think he is arguing from the point of view of the stress/strain difference between the gas and the metal. The metal's stress/strain ratio is much higher than the gases, so it takes most of the load.

John

Offline cambrianera

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #162 on: 09/18/2017 12:58 PM »
I think he is arguing from the point of view of the stress/strain difference between the gas and the metal. The metal's stress/strain ratio is much higher than the gases, so it takes most of the load.

John
Well, compressed gas actively pushes against the bulkheads (pressure x area, should be seen as an already compressed spring); metal takes only the remainder of the load (or stand in tension, depending on flight phase and vertical position considered).
Oh to be young again. . .

Offline livingjw

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #163 on: 09/18/2017 04:06 PM »
I think he is arguing from the point of view of the stress/strain difference between the gas and the metal. The metal's stress/strain ratio is much higher than the gases, so it takes most of the load.

John
Well, compressed gas actively pushes against the bulkheads (pressure x area, should be seen as an already compressed spring); metal takes only the remainder of the load (or stand in tension, depending on flight phase and vertical position considered).
  you are right. Duh!
« Last Edit: 09/18/2017 04:09 PM by livingjw »

Offline hkultala

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Re: F9 - S2 reusable modification as evolution steps to BFS(ITS)
« Reply #164 on: 09/19/2017 07:34 PM »
Is there any sense in having the f9-heavy central core stage and the upper-stage the same length as below (modified from image on Wikipedia):


The core stage would be a Falcon-5 with legs and the upper-stage a falcon-3 (1 standard merlin + 2 twin-chamber over-expanded merlins).

So many different new stages and even a new engine type for minimal benefit. Costs would skyrocket, absolutely no sense.

What is the point of the hyphothetical "twin-chamber over-expanded engine"? That is the part that adds the "absolutely" part into my judgement.

Quote

This might allow attempting cradle landings of the upper-stage.

No, it would not make it any easier, would only make it harder.


Yes there are new stages. The Falcon 5 is a shortened Falcon-9 with 4 engines removed. The upper-stage is then based on the same tankage.

.. which would mean quite badly balanced stages. The core stage might be somewhat underpowered after booster separation, and upper stage would be too big for BLEO payloads, as too much mass would have to go to the final orbit. This would hurt payload considerably.

Quote
The upper-stage is much thinner/finer than the current one and might be more stable re-entering.  It would re-enter base first, be quite fluffy and maybe better shaped for hypersonic flight than a more stumpy stage.

Then it would simply burn when re-entering. The engines cannot withstand the temperature of re-entry without
heat shield. And structurally it would also be quite weak, easily breaking apart fro aerodynamic forces. (speed on re-entry is much higher than any speed current F9 1st stage has to fly in atmosphere.

Quote
It would have more rotational inertia. It is easier to balance a broom on your hand than a pencil. The US has a central sea-level engine that can be used for landing/cradling. 

You are solving the wrong problem, and making the real problems much worse.

Quote
The twin chamber engines have better isp than the regular merlins for acceptable upper-stage economy. 2 x fat-nozzled twin-chamber engines  + 1 regular engine allow 3 engines to pack better under a stage. The four fat nozzles arranged in an exploded square surrounding the smaller sea-level engine.

No, twin chambers do not help at all for isp.

Quote
Do we know how robust the current vacuum merlin nozzle would be through re-entry?

We know that it cannot survive reentry from stage speed because stage 1 has to do the re-entry burn to make it survive. Stage 2 has many times higher kinetic energy/mass than stage 1.

Quote
The combined stack will be shorter (FH is quite thin/fine.)

.. and you are removing fuel and tankage from the part where it's needed the most - core stage

Quote
Would you prefer two stacked Falcon-3-s for more commonality, or a Falcon-5 with a stretched conventional upper-stage

falcon 5 with slightly streeched upper stage is the least bad option.

two 3's would be very badly unbalanced design.

Quote
(new tangent: would a four chambered vacuum merlin - like the rd-0124 - be more robust through re-entry?)

Even more stupid idea.

Forget the multi-chamber engines. They make no sense.

Russians use them because they wanted to use single turbopump, and couldn't make big chamber due combustion instability problems, and to avoid too big nozzles that hurt T/W. Neither of these reasons apply for any falcon-derivate rocket.

SpaceX wants to have the turbopumps they have and mass-produce those, developing twice bigger to run two chambers would make absolutely no sense at all. Reliability would suffer(worse engine-out options), manufacturing costs would skyrocket as less mass manufacturing, need to develop new parts for no gain at all. NO SENSE AT ALL.

« Last Edit: 09/20/2017 09:03 AM by hkultala »

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