Author Topic: Operating the ITS spaceship or tanker as an SSTO launch vehicle.  (Read 18732 times)

Online envy887

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Here are the engines from the full vehicle cutaway image, overlaid next to each other. There is no perspective distortion, as you can tell from the 12m main tank diameters matching perfectly. The "sea level" nozzles on the spaceship are clearly more expanded than those on the booster. Measuring the image gives a value of about 20% larger.

Perhaps this is meaningless overanalysis, but I don't think so. And it does make sense to have larger nozzles on the spaceship, for reasons noted above.

Offline OneSpeed

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Yes, by comparing the IAC Mexico images with those generated using a simple rendering program:

Raptor 40, 1.51m
Raptor 50, 1.78m
Raptor Vac, 3.8m
« Last Edit: 01/13/2017 09:51 PM by OneSpeed »

Offline dglow

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Good to see! I pointed this out back in October but nobody took the bait.  ???

Offline RocketmanUS

Are the images to scale?

Slide 36 from the SX PDF on ITS
Quote:
" 3 Sea-Level - 361s ISP "
" 6 Vacuum - 382s ISP "

When it says 3 Sea Level that is referring to three sea level engines. I believe those are the vacuum ratings for both engines.

Slide 26 does show the BFS sea level engine bell larger than the BFR sea level engine. However this image might not be to scale. Compare the length and diameter of the BFS and the BFR.

We would need to hear from SpaceX on this.

Anyway Raptor 80 could possible give the best mass to orbit if it's specs are correct.
Sea level thrust and ISP?
Vacuum thrust and ISP?
Chamber pressure?
Dry mass.
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Online envy887

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Are the images to scale?

Slide 36 from the SX PDF on ITS
Quote:
" 3 Sea-Level - 361s ISP "
" 6 Vacuum - 382s ISP "

When it says 3 Sea Level that is referring to three sea level engines. I believe those are the vacuum ratings for both engines.

Slide 26 does show the BFS sea level engine bell larger than the BFR sea level engine. However this image might not be to scale. Compare the length and diameter of the BFS and the BFR.

They are to scale in the image I posted: the tank diameters are identical and there is no perspective distortion.

Quote
We would need to hear from SpaceX on this.

Anyway Raptor 80 could possible give the best mass to orbit if it's specs are correct.
Sea level thrust and ISP?
Vacuum thrust and ISP?
Chamber pressure?
Dry mass.

From a RPA sim using LOX/LCH4 at 300 bar, 3.800 O/F, 80:1 expansion from .0582 m2 throat, 100% bell nozzle length and 99.50% reaction efficiency, the expected actual performance (not theoretical best performance, which is a couple percent higher) is:
SL: 3005 kN at 323.2 s
Vacuum: 3479 kN at 372.2 s

My mass estimate is ~2800 kg
. Mvac sans nozzle is ~600 kg at 934 kN, and mass scales roughly with thrust, which would put the R80 at 2230 kg. I'm adding ~570 kg for the 80:1 nozzle and added mass penalties of FFSC cycle compared to GG and methalox compared to kerolox.

Edit: the nozzle exit diameter for this specific configuration would be 2.43 meters. The throat diameter is the same as the 200:1 engine.
« Last Edit: 01/14/2017 01:34 PM by envy887 »

Offline john smith 19

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This is highly entertaining but I've yet to see any definition on how this stage will reenter and land the key issues.

Isp is obviously important, as is how it varies on altitude. So is heating rate as that will set the amount of TPS you have to carry (or fuel you stream through the bells in the early stages of reentry).

It's known that all Vertical Takeoff & Landing SSTO concepts are critically dependent on Isp and structural mass fraction.

The SSME bells were IIRC 77:1 but Raptor should be able to do better. That matters if you're tracking Isp change with altitude and you plan to use a straight single position Rao nozzle.
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Online envy887

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I don't think any modifications have been proposed that would significantly affect the EDL sequence as described in the IAC presentation. EDL should work essentially the same for SSTO ITS as TSTO ITS.

Offline Robotbeat

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It's known that all Vertical Takeoff & Landing SSTO concepts are critically dependent on Isp and structural mass fraction.
...
This is true whether vertical, horizontal, or diagonal takeoff/landing SSTO RLV.

Even true for Skylon, which has to have ridiculously* good mass fraction (from a volumetric perspective).


*but perhaps still achievable
« Last Edit: 01/16/2017 05:03 PM by Chris Bergin »
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Offline OneSpeed

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...
We would need to hear from SpaceX on this.
...

Here are the Raptor figures I'm using in SpaceSim. Most of these figures have been published by SpaceX or mentioned in the Reddit AMA. I've simply interpolated to get my Raptor 80 figures. The mass estimates are envy's, they sound pretty reasonable to me.

   Cycle            Full-flow staged combustion
   Oxidiser         Subcooled liquid oxygen
   Fuel            Subcooled liquid methane
   Chamber pressure   300 bar or 30 MPa
   Throttle capability 20% to 100% thrust

   Sea-level Nozzle
      Expansion Ratio   40
      Thrust (SL)      2,842 kN
      Thrust (Vac)   3,061 kN
      Isp (SL)      333s
      Isp (Vac)      360s
      Diameter      1.51m
      Mass         2400 kg
      
      Expansion Ratio   50
      Thrust (SL)      3,094 kN
      Thrust (Vac)   3,333 kN
      Isp (SL)      334s
      Isp (Vac)      361s
      Diameter      1.78m
      Mass         2500 kg
      
      Expansion Ratio   80
      Thrust (SL)      3,200 kN
      Thrust (Vac)   3,400 kN
      Isp (SL)      336s
      Isp (Vac)      368s
      Diameter      2.4m
      Mass         2800 kg
      
   Vacuum Nozzle
      Expansion Ratio   200
      Thrust         3,500 kN
      Isp            382s
      Diameter      3.8m
      Mass          3000 kg

Offline RocketmanUS

...
We would need to hear from SpaceX on this.
...

Here are the Raptor figures I'm using in SpaceSim. Most of these figures have been published by SpaceX or mentioned in the Reddit AMA. I've simply interpolated to get my Raptor 80 figures. The mass estimates are envy's, they sound pretty reasonable to me.

   Cycle            Full-flow staged combustion
   Oxidiser         Subcooled liquid oxygen
   Fuel            Subcooled liquid methane
   Chamber pressure   300 bar or 30 MPa
   Throttle capability 20% to 100% thrust

   Sea-level Nozzle
      Expansion Ratio   40
      Thrust (SL)      2,842 kN
      Thrust (Vac)   3,061 kN
      Isp (SL)      333s
      Isp (Vac)      360s
      Diameter      1.51m
      Mass         2400 kg
      
      Expansion Ratio   50
      Thrust (SL)      3,094 kN
      Thrust (Vac)   3,333 kN
      Isp (SL)      334s
      Isp (Vac)      361s
      Diameter      1.78m
      Mass         2500 kg
      
      Expansion Ratio   80
      Thrust (SL)      3,200 kN
      Thrust (Vac)   3,400 kN
      Isp (SL)      336s
      Isp (Vac)      368s
      Diameter      2.4m
      Mass         2800 kg
      
   Vacuum Nozzle
      Expansion Ratio   200
      Thrust         3,500 kN
      Isp            382s
      Diameter      3.8m
      Mass          3000 kg
First thanks  :) .

First we have this quote-
Questions and Answers from the I am Elon Musk, ask me anything about becoming a spacefaring civ! AMA on the /r/spacex Reddit site pertaining to reentry and landing:


https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94vcmw/?context=3
Quote
Question from __Rocket__

ITS Spaceship design question II.:

The ITS Spaceship has two mystical spherical tanks, marked green in this slightly edited image. The whole tank design looks very exciting, and there's rampant speculation on this sub about the purpose of those spherical tanks:

    are they for landing fuel?
    ... or are they storing 'hot' gaseous propellants as part of the autogenous propellant pressurization system?
    ... or are they used for on-orbit propellant densification to store vapor before it's liquefied again?

All of the above perhaps?



Answer from /u/ElonMuskOfficial

Those are the header tanks that contain the landing propellant. They are separate in order to have greater insulation and minimize boil-off, avoid sloshing on entry and not have to press up the whole main tank.


https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94vdk1/?context=3
Quote
Question from /u/_rocketboy

Also, why does the booster only have one in 1 tank?



Answer from /u/ElonMuskOfficial

The liquid oxygen transfer tube serves as the header tank for ox

https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94v8p8/?context=3
Quote
Question from /u/FoxhoundBat

    Overall is the landing architecture of ITS booster and distances needed to be covered to be same as Falcon 9s? Boostback, re-entry burn, landing burn?

    Could you give us nuggets on what changes the final Falcon 9 version (v1.3) you mentioned will have? Uprated engines obviously from 170k to 190k lbf, but what else? Is it mostly geared towards reusabilty over performance?

    Gwynne mentioned 2 weeks ago that F9 v1.2 will be reused only once or twice while v1.3 should be reused up to 10 times. Can you talk about what are the limiting factors for Falcon 9 reuse?



Answer from /u/ElonMuskOfficial
The big booster will have an easier time of things than Falcon, as the mass ratio of the stages is lower and it will have lower density. Net result is that it won't come in quite as hot and fast as Falcon, so Falcon should be a bounding case on the big booster.

Final Falcon 9 has a lot of minor refinements that collectively are important, but uprated thrust and improved legs are the most significant.

Actually, I think the F9 boosters could be used almost indefinitely, so long as there is scheduled maintenance and careful inspections. Falcon 9 Block 5 -- the final version in the series -- is the one that has the most performance and is designed for easy reuse, so it just makes sense to focus on that long term and retire the earlier versions. Block 5 starts production in about 3 months and initial flight is in 6 to 8 months, so there isn't much point in ground testing Block 3 or 4 much beyond a few reflights.

https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94ukv8/?context=3
Quote
Question from /u/termderd

We got a pretty good idea of what a Mars EDL looks like, but can you explain how the ITS and the Tanker plan to do an Earth EDL? Having talked with you at IAC about the Mars entry, we learned that there's very powerful thrusters that can handle attitude control. These work great for the Martian atmosphere, but what about on earth? There doesn't appear to be grid fins and the thrusters obviously have less authority here on earth, so what's the trick?

Thanks for your time!

    Tim Dodd, The Everyday Astronaut



Answer from /u/ElonMuskOfficial

Good question -- that wasn't shown at IAC. The spaceship and tanker would have split body flaps for pitch and roll. Probably just use the attitude control thrusters for yaw.

https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94ub7h/?context=3
Quote
Question from /u/Tesla_X_City:
If I recall correctly on one of the slides it mentioned that there it will be 4-6 G's upon reentry. It does not specify, however, whether that will be during the landing burn or aerobreaking. It would be nice if that is clarified as well.



Answer from /u/ElonMuskOfficial

The spaceship would be limited to around 5 g's nominal, but able to take peak loads 2 to 3 times higher without breaking up.

Booster would be nominal of 20 and maybe 30 to 40 without breaking up.

https://www.reddit.com/r/spacex/comments/590wi9/i_am_elon_musk_ask_me_anything_about_becoming_a/d94u6zk/?context=3
Quote
Question from /u/TheVehicleDestroyer

Hi Elon. Ive got 3 questions on the ITS vehicle specs:

    Can you divulge what the Vacuum Thrust+Isp figures are for the Sea-Level Raptor variant?

    The ITS booster is able to hover. Will it ever use this capability to better ensure a successful landing at the expense of some small gravity losses, or is it hoverslams all the way?

    What is the expected maximum acceleration that the ITS booster can withstand during entry/landing?

Thanks for everything.



Answer from /u/ElonMuskOfficial

    Approx 360 sec vacuum Isp and 290 metric tons of thrust
    A high acceleration landing is a lot more efficient, so there wouldn't be any hovering unless it encountered a problem or unexpected wind conditions. A rocket that lands slowly is wasting a lot of fuel.
    Aiming for 20 g's


Then we have this quote-
Just noticed that SpaceX posted higher resolution photos of the raptor test fire on flickr than were attached to Elon's original tweets (as originally posted below). I can't see these higher resolutions posted earlier in this, or the previous ITS propulsion thread.

Quote from: Elmar Moelzer link=topic=34197.msg1588736#msg1588736
Elon Musk on Twitter:
SpaceX propulsion just achieved first firing of the Raptor interplanetary transport engine
https://twitter.com/elonmusk/status/780280440401764353

Production Raptor goal is specific impulse of 382 seconds and thrust of 3 MN (~310 metric tons) at 300 bar
https://twitter.com/elonmusk/status/780275236922994688
So what do you make of this?

As the Raptor 80 could give the best payload mass should we use it for the BFS SSTO ( tanker version modified to carry payload at it's top portion )? And double check the specs for this Raptor variant?
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Offline RocketmanUS

For a minimal design change version, how would a vehicle with only six Raptor 80 work, leaving the 3 central sea level Raptor unchanged? That would work without changing plumbing and thrust structure for maximum commonality with the standard ITS.

I've run a quick simulation of your suggestion, and the results look like this:

ConfigurationMass to 300 x 300Vehicle massPayload
9 x Raptor 400 mT89 mT0mT
9 x Raptor 50190 mT90 mT28mT
3 x Raptor 50 + 6 x Raptor 80196 mT92 mT32mT
12 x Raptor 80240 mT100 mT67mT

So, perhaps an extra 4mT to orbit, but you would have the extra cost of development of the Raptor 80.
OK , so I just subtracted the six vacuum Raptor engine thrust from the BFS total vacuum thrust and divided by three to get the vacuum thrust of the sea level engines. 333,333,333 N, so that is the same for your figures for the Raptor 50 , so will have to call it three versions for Raptor.

9 Raptor 50 at ~28 mt would be impressive.
12 Raptor 80 at ~67 mt would be even more impressive.
That is even after adding in the mass for a hatch ( or door ) along with a cargo bay the up/down payload mass.

What is the propellant mass at liftoff for the 9 and 12 Raptor 80 models?

Has anyone measured the volume of the BFS crew version propellant tanks? If so how much propellant can they hold by mass? How high will the tanks rise into the cargo area for these concept SSTO BFR's ( 9 and 12 engine versions ). This is to calculate how much possible volume is left for possible payload ( height of tank to top of BFS ).
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Offline john smith 19

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Since expansion ratio is quite important to this it might help to know what at what level flow separation is likely to take place.

A common rule of thumb is when the nozzle outlet pressure is < 0.4 of the ambient pressure. More accurate tests based on real nozzles reckon this is conservative and 0.34 of Pambient is safe.

At 300 bar with 1 Atm at 101326 Pa and 14.7psi to 1 Atm that's 4352 psi, close to double the SSME's chamber pressure.

That's suggests you could go higher than 80 expansion ratio at Earth SL takeoff, which is the critical limit for takeoff thrust, stresses etc.  Bigger expansion ratio is the #1 choice of engine designers when they want more Isp, but I don't have a feel for how chamber pressure scales with expansion ratio.

If you're going to chase the VTOL SSTO concept you need all the low hanging fruit you can get to make it work.

"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 OneSpeed

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Since expansion ratio is quite important to this it might help to know what at what level flow separation is likely to take place.

A common rule of thumb is when the nozzle outlet pressure is < 0.4 of the ambient pressure. More accurate tests based on real nozzles reckon this is conservative and 0.34 of Pambient is safe.

At 300 bar with 1 Atm at 101326 Pa and 14.7psi to 1 Atm that's 4352 psi, close to double the SSME's chamber pressure.

That's suggests you could go higher than 80 expansion ratio at Earth SL takeoff, which is the critical limit for takeoff thrust, stresses etc.  Bigger expansion ratio is the #1 choice of engine designers when they want more Isp, but I don't have a feel for how chamber pressure scales with expansion ratio.

If you're going to chase the VTOL SSTO concept you need all the low hanging fruit you can get to make it work.

I agree that given the outstanding chamber pressure of the Raptor, the expansion ratio could be a little higher, although you would loose some of the ability to throttle deeply. The sims I've run so far indicate that going from 9 Raptor 50s to 6 Raptor 80s + 3 Raptor 50s gains you about 4mT to orbit. Going to say 9 Raptor 90s might gain you another couple of tons.

However, for SSTO, increasing thrust made a far bigger difference. Having 12 Raptor 80s added 39mT to the payload, mostly because of reduced gravity losses, rather than better Isp. I would argue that T/W is actually the low hanging fruit for SSTO of the ITS (not saying better Isp doesn't help).

Offline guckyfan

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However, for SSTO, increasing thrust made a far bigger difference. Having 12 Raptor 80s added 39mT to the payload, mostly because of reduced gravity losses, rather than better Isp. I would argue that T/W is actually the low hanging fruit for SSTO of the ITS (not saying better Isp doesn't help).

But going to 12 Raptor would require a basically new thrust structure, almost a new rocket. Staying at 9 would be much less development intensive so a lot cheaper.

Online Semmel

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However, for SSTO, increasing thrust made a far bigger difference. Having 12 Raptor 80s added 39mT to the payload, mostly because of reduced gravity losses, rather than better Isp. I would argue that T/W is actually the low hanging fruit for SSTO of the ITS (not saying better Isp doesn't help).

But going to 12 Raptor would require a basically new thrust structure, almost a new rocket. Staying at 9 would be much less development intensive so a lot cheaper.

So what this basically means is.. if Raptor takes a similar development as the Merlin and increases its thrust by, say 50% over the first couple of years of using it, an ITS SSTO might be a reasonable thing to look for. I wouldnt count on that though since ITS might grow with the Raptor performance to maximize payload to Orbit.

Given your simulations OneSpeed, I dont think we will see an ITS SSTO and I agree with Guckyfan, a 12-Raptor version of the ITS is probably not going to happen. Time will tell of course.

Online envy887

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Increasing expansion does eventually lead to seperation at sea level, at about 140:1, but SL thrust drops to 2715 kN per engine and ISP drops to 292s. On a thrust-limited design this is a poor trade-off, even though vacuum ISP is close to 380.

However, instead of nine 80:1 nozzles it might make sense to have a mix of 50:1 and 140:1. The former have much higher thrust and ISP at SL, and the latter have higher thrust and ISP in vacuum, but both are operable (if needed) across the entire flight. The 50:1 engines could be shut down to improve overall ISP once the vehicle is high enough (and light enough) that the 140:1 nozzles can maintain optimal acceleration.

Offline OneSpeed

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Increasing expansion does eventually lead to seperation at sea level, at about 140:1, but SL thrust drops to 2715 kN per engine and ISP drops to 292s. On a thrust-limited design this is a poor trade-off, even though vacuum ISP is close to 380.

However, instead of nine 80:1 nozzles it might make sense to have a mix of 50:1 and 140:1. The former have much higher thrust and ISP at SL, and the latter have higher thrust and ISP in vacuum, but both are operable (if needed) across the entire flight. The 50:1 engines could be shut down to improve overall ISP once the vehicle is high enough (and light enough) that the 140:1 nozzles can maintain optimal acceleration.

That's an interesting concept, but for SL that gives:

(3 * 3,094 + 6 * 2,715) / 9.8 = 2609mT thrust.

So T/W = 2609 / 2590 = 1.01, the same as for 9 Raptor 40s. I suspect the gravity losses at low altitude would be similar, with a similar payload to a 300 x 300 km orbit (unfortunately none).

...
Given your simulations OneSpeed, I dont think we will see an ITS SSTO and I agree with Guckyfan, a 12-Raptor version of the ITS is probably not going to happen. Time will tell of course.

Actually I suspect the same, but I was asked to provide the sims, so I'm just trying to help where I can. More thrust is the only way I can see the SSTO having a chance, but even then, it seems more efficient to use the full BFR/BFS stack, and if small payloads are required, deliver more of them with each flight.

Offline john smith 19

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Actually I suspect the same, but I was asked to provide the sims, so I'm just trying to help where I can. More thrust is the only way I can see the SSTO having a chance, but even then, it seems more efficient to use the full BFR/BFS stack, and if small payloads are required, deliver more of them with each flight.
This has always been the issue with VTO SSTO.

VTO TSTO gives you 3-4% of GTOW. Historically people have been willing to accept 1% of GTOW for SSTO.  There is only one known architecture that can deliver a TSTO payload and that can't land on Mars.
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Offline Robotbeat

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I fail to see how vertical or horizontal makes any difference there. The problem is the same.
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Offline RocketmanUS

For a minimal design change version, how would a vehicle with only six Raptor 80 work, leaving the 3 central sea level Raptor unchanged? That would work without changing plumbing and thrust structure for maximum commonality with the standard ITS.

I've run a quick simulation of your suggestion, and the results look like this:

ConfigurationMass to 300 x 300Vehicle massPayload
9 x Raptor 400 mT89 mT0mT
9 x Raptor 50190 mT90 mT28mT
3 x Raptor 50 + 6 x Raptor 80196 mT92 mT32mT
12 x Raptor 80240 mT100 mT67mT

So, perhaps an extra 4mT to orbit, but you would have the extra cost of development of the Raptor 80.
12 Raptor 80 how much propellant mass were you using?
What is the maximum propellant mass for the tanker for their propellant tank size?
SX PDF slide 36 show 2,500 mt propellant.
And shows 380 mt propellant delivered to orbit. I assume that was part of the 2,500 mt and not extra.
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