Author Topic: ITS Propulsion – The evolution of the SpaceX Raptor engine  (Read 61697 times)

Offline baldusi

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #100 on: 11/08/2016 08:30 PM »
I'm super excited. But as you said no info to work with.
I think its not only the dearth of info, but the high inconsistency on the available one. I had more than 10 questions regarding Raptor in the Reddit AMA for Elon, but obviously none was answered. Apparently "how do you feel ..." are a lot more interesting than the use of expander cycle for the low pressure turbopump.

Expander cycle for the low pressure turbopump???
Its a relatively common trick. I didn't saw anything like that in the picture, just a speculative question. But it is a trick used by the SSME. They use a low pressure pump to avoid cavitation. And run it from the supercritical fuel that's output by the regen cooling loop.
KBKhA RD-0162/SD use the expander cycle to run the mail fuel pump. And that was a 2MN engine. So there is some significant power availability from the expander cycle for a 3MN rocket. A pity not to use it.

Offline jpo234

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #101 on: 01/09/2017 02:10 PM »
Are there any updates about Raptor development after the September test?

Offline philw1776

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #102 on: 01/10/2017 07:30 PM »
Are there any updates about Raptor development after the September test?

None about development or test
“When it looks more like an alien dreadnought, that’s when you know you’ve won.”

Online FutureSpaceTourist

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #103 on: 01/13/2017 10:37 AM »
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
« Last Edit: 01/13/2017 10:40 AM by FutureSpaceTourist »

Offline rockets4life97

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #104 on: 01/27/2017 02:56 AM »
Any word on more tests? Anybody have a guess about how long they will test this initial engine before moving to an upgraded version?

Offline envy887

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #105 on: 03/10/2017 08:21 PM »
Here is a strictly hypothetical question.

Assuming this 1,000 kN demonstrator reaches a 30 MPa operating chamber pressure, how big/wide would a 50:1 ratio nozzle be for it? Moreover, what would be the most effective/efficient nozzle ratio that it could have, assuming it is used for first stage propulsion (among 8 other engines) and slow/low S1 separation for RTLS duties?

For a booster engine 50:1 would be about right. Cycle is the same so the nozzle scales with area so its diameter scales with the square root of the thrust ratio:
                                          40:1 diam = 1.7 x sqrt(1 / 3.05)         = . 97 m  (~38 inches).
                                          50:1 diam = 1.7 x sqrt(50 / 40 /3.05) = 1.09 n (~ 43 inches)

Many thanks for that. A couple more questions to anyone interested to answer (again, this is a hypothetical scenario).

What is the diameter of the current M9 nozzle?
If we assume that the material, width and height of the current F9 S1 remains constant, and that the common bulkhead is moved to adjust, given:

1. The known propellant ratio for the Raptor Demonstrator.
2. An SL thrust of 870kN and Vac thrust of 930kN.
3. An SL Isp of 330s and Vac Isp of 358s
4. A dry stage weight of 27 metric tons.

What would the performance delta be against the current F9 S1?

I'm not asking whether something like this is possible, probable, practicable or wanted/needed. Just want to understand the comparative difference between one engine and the other in a hypothetical scenario. I assume that the difference would be rather small, both due to having less propellant on the stage and Isp not being the most important factor in the two re-usable scenarios that F9 S1 covers (RTLS and DPL S1-S2 separations).

Adjusting for lower propellant density and assuming similar engine TWR, the high pressure FFSC methalox still gets 31% more payload to LEO and 38% more payload to GTO compared to low pressure GG kerolox:

Using http://www.silverbirdastronautics.com/LVperform.html
To 185 km x 28.5 deg circular LEO with no fairing and 0.5% residuals:

21162 kg for kerolox S1: 24000 kg dry, 430000 kg prop, 8000 kN avg, 297 sec avg; S2: 4500, 115000, 934, 348.

25743 kg for methalox S1: 24000 kg dry, 360000 kg prop, 8100 kN avg, 348 sec avg; S2: 4500, 96000, 1000, 374.

To 185 x 38500 km x 28.5 deg GTO with 4000 kg fairing discarded at 220 sec, and 0.5% residuals:

7006 kg for kerolox S1: 24000 kg dry, 430000 kg prop, 8000 kN avg, 297 sec avg; S2: 4500, 115000, 934, 348.

9680 kg for methalox S1: 24000 kg dry, 360000 kg prop, 8100 kN avg, 348 sec avg; S2: 4500, 96000, 1000, 374.

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #106 on: 03/14/2017 06:17 PM »
I've run numbers on RPA-lite to calculate a family of Raptor engines differing only in Expansion Ratio (ER). I have calibrated my model considering only two authoritative sources: The vacuum numbers for Raptor in the IAC lecture and the stage drawings that are supposedly directly from CAD. This means I'm using 3.7 O/F from spaceflight101 tank measurements instead of the 3.8 O/F that Elon said. I attached my RPA-lite configuration file for the Raptor 200 (just change the extension to .cfg). For the others I only variated the ER.

I used the 'freezing at area ratio' to aim precisely at 382 s isp for the Raptor 200. It gave an pretty high number of 12 and is still undershooting the SL variants of the engine. The RPA guys use 6 for their RD-253 (N2O4/UDMH) performance validation and still undershot the ISP too, especially at SL. So maybe more is adequate for a methane raptor, I don't know. It is set lower for other fuel types and R7 found that 3 is adequate to simulate a Russian methane rocket engine.

Leaving the throat diameter fixed at 0.2685m and using the measurements from OneSpeed, by simple scaling I get an ER for booster engines of 32:1 and 44:1 for the BFS SL engines. I assume that the 3050 kN 334s at 40 ER SL engine described in the IAC slides is in fact the Raptor 32 while the 361 s vacuum isp is the Raptor 44. RPA has undershot both slightly. The Raptor 40 is as far as I understand just a middle of the way designation to talk about the performance of an average SL Raptor, but I ran numbers for it too anyway, as well as the usually discussed Raptor 50.

I also ran numbers for other intermediate ER, for the benefit of those who are dreaming with a BFS SSTO (me included). 116:1 being one that fits 9 in the perimeter of BFS and 130:1 being the maximum ER that RPA-lite doesn't warns me against flow separation at SL. Some altitude performance analysis graphs are attached too. They seem based on Theoretical performance, not the estimated delivered performance.


    Nozzle size    |        Sea Level      |          Vacuum       | Optimal Expansion |
 ER | Diameter (m) | Thrust (kN) | Isp (s) | Thrust (kN) | Isp (s) |  H (km) |  P (atm)|
----|-------------:|------------:|--------:|------------:|--------:|--------:|--------:|
 32 |     1.52     |    3044     |  332.0  |    3234     |  353.0  |   0.00  |   1.002 |
 40 |     1.70     |    3037     |  331.5  |    3274     |  357.4  |   2.33  |   0.753 |
 44 |     1.78     |    3029     |  330.7  |    3290     |  359.1  |   3.29  |   0.667 |
 50 |     1.90     |    3015     |  329.1  |    3311     |  361.4  |   4.55  |   0.566 |
 57 |     2.03     |    2994     |  326.8  |    3332     |  363.7  |   5.79  |   0.479 |
 80 |     2.40     |    2908     |  317.4  |    3382     |  369.2  |   8.82  |   0.312 |
116 |     2.89     |    2746     |  299.8  |    3434     |  374.8  |  11.88  |   0.195 |
130 |     3.06     |    2678     |  292.3  |    3448     |  376.4  |  12.80  |   0.169 |
200 |     3.80     |    2315     |  253.0  |    3500     |  382.0  |  16.27  |   0.098 |



I don't know how to force a fixed width font in this forum (edit: now I know, thanks). The results are inside RPA error margin, especially considering that it should not be as tuned for methane because the lack of real engine data to check against.

I'm ignoring completely this latest information on Raptor, as it suggests a smaller engine with a vacuum thrust at 200:1 ER in the 3125 kN range, while the IAC slides said 3500 kN.
« Last Edit: 03/16/2017 12:09 AM by Manabu »

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #107 on: 03/14/2017 06:24 PM »
I also did a throttling analysis on the same basis. The Raptor 40 isn't quite capable of throttling down to 20% before flow separation at SL, according to RPA-lite. But with 32:1 ER it can, and with 44:1 it can throttle down to about 30%. Maybe some nozzle tricks may prove those numbers too conservative.

Offline envy887

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #108 on: 03/14/2017 06:46 PM »
I also did a throttling analysis on the same basis. The Raptor 40 isn't quite capable of throttling down to 20% before flow separation at SL, according to RPA-lite. But with 32:1 ER it can, and with 44:1 it can throttle down to about 30%. Maybe some nozzle tricks may prove those numbers too conservative.

Nice work!

For the throttled engines, are you plotting chamber pressure ratios or thrust ratios? Because of atmospheric back-pressure at sea level slowing the exhaust, throttling the chamber pressure to 20% will produce less than 20% thrust.

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #109 on: 03/14/2017 07:47 PM »
Nice work!

For the throttled engines, are you plotting chamber pressure ratios or thrust ratios? Because of atmospheric back-pressure at sea level slowing the exhaust, throttling the chamber pressure to 20% will produce less than 20% thrust.
I specified the interval as thrust ratios, where 1.0 corresponds to the nominal thrust. RPA-lite did the rest for me. But good observation, I haven't thought about that.

EDIT: Another thing to have in mind is that those numbers use the SL performance that I estimated with RPA-lite, that is a bit lower than the ones confusingly said by SpaceX. I'm also using the 3.7 O/F that gives a little less thrust for a given ISP.

I redid the Throttled chamber performance analysis with a more orthodox 3.8 O/F, pure shifting equilibrium model for the nozzle and reaction efficiency manually raised to 99.4 to match the Raptor 40 IAC numbers. Graph in the attachment and here the engine parameters compared to the ones in the other table:


      Nozzle size     |        Sea Level      |          Vacuum       | Optimal Expansion |
  ER   | Diameter (m) | Thrust (kN) | Isp (s) | Thrust (kN) | Isp (s) |  H (km) |  P (atm)|
-------|-------------:|------------:|--------:|------------:|--------:|--------:|--------:|
40     |     1.70     |    3037     |  331.5  |    3274     |  357.4  |   2.33  |   0.753 |
40  V2 |     1.70     |    3052     |  334.1  |    3287     |  359.8  |   1.69  |   0.815 |
200    |     3.80     |    2315     |  253.0  |    3500     |  382.0  |  16.27  |   0.098 |
200 V2 |     3.80     |    2361     |  258.4  |    3536     |  387.0  |  15.55  |   0.110 |


In the end the only thing that seems to have affected the plot is the O/F ratio, and then only a little, as RPA-lite seems to also use Theoretical performance instead of the estimated delivered performance in this plot.
« Last Edit: 03/16/2017 12:12 AM by Manabu »

Offline OneSpeed

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #110 on: 03/15/2017 07:58 PM »
I don't know how to force a fixed width font in this forum.

You can create a table in the reply editor, using the table tags, but it is a bit laborious:

Nozzle sizeSea LevelVacuum-OptimalExpansion
ERDiameter (m)Thrust (kN)Isp (s)Thrust (kN)Isp (s)H (km)P (atm)
401.703037331.53274357.42.330.753
40  V21.703052334.13287359.81.690.815
2003.802315253.03500382.016.270.098
200 V23.802361258.43536387.015.550.110

Is that what you are after?

Offline nacnud

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #111 on: 03/15/2017 08:03 PM »
This may help in the future, but test it first!

http://www.teamopolis.com/tools/bbcode-table-generator.aspx

Online AnalogMan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #112 on: 03/15/2017 08:42 PM »
I don't know how to force a fixed width font in this forum.

You can create a table in the reply editor, using the table tags, but it is a bit laborious:

Nozzle sizeSea LevelVacuum-OptimalExpansion
ERDiameter (m)Thrust (kN)Isp (s)Thrust (kN)Isp (s)H (km)P (atm)
401.703037331.53274357.42.330.753
40  V21.703052334.13287359.81.690.815
2003.802315253.03500382.016.270.098
200 V23.802361258.43536387.015.550.110

Is that what you are after?

You can force a fixed pitch font using the
[tt] and [/tt]
tags.  If using the simple forum editor in preview mode then you can also highlight the relevant text and click the "Tt" button - this inserts the tags for you.

This produces a monospaced teletype font - this is what it looks like applied to the text your table:

    Nozzle size    |        Sea Level      |          Vacuum       | Optimal Expansion |
 ER | Diameter (m) | Thrust (kN) | Isp (s) | Thrust (kN) | Isp (s) |  H (km) |  P (atm)|
----|-------------:|------------:|--------:|------------:|--------:|--------:|--------:|
 40 |     1.70     |    3037     |  331.5  |    3274     |  357.4  |   2.33  |   0.753 |
40  V2 |  1.70     |    3052     |  334.1  |    3287     |  359.8  |   1.69  |   0.815 |
200 |     3.80     |    2315     |  253.0  |    3500     |  382.0  |  16.27  |   0.098 |
200 V2 |  3.80     |    2361     |  258.4  |    3536     |  387.0  |  15.55  |   0.110 |
« Last Edit: 03/15/2017 08:43 PM by AnalogMan »

Offline Manabu

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #113 on: 03/16/2017 12:33 AM »
Thanks all above, I fixed the tables using the AnalogMan advice. The bbcode table is laborious to make, even with that website, while I already have a workflow for those fixed width tables and they are more "portable". But maybe I can use for some future tables to make them a little prettier.

I found a small problem in my simulation. When I went to look the logs by curiosity, I found this silent warning:
Quote
WARNING: Temperature T=93.00 K could not be assigned to the species "CH4(L)". Using T=298.15 K instead.
The minimum temperature supported for CH4 is 100 K, and that reduces the isp compared to 298.15 K by about 2 s, all else the same. When increasing the freezing area ratio to match the 382 Raptor 200 isp, the Raptor 32 isp drop up to 2 s compared to the previous simulation.

But I'm right in using those sub-cooled temperatures as they are in the tanks? Or should I use high temperatures and pressures for the fuel (and maybe the oxidizer too) because the engine is regenerative cooled? This would reduce a little, but not eliminate, the gap between SpaceX stated SL performance and my RPA-lite simulations.
« Last Edit: 03/16/2017 12:37 AM by Manabu »

Offline spacenut

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #114 on: 03/21/2017 09:09 PM »
How far along is the Raptor engine?  Any word as to when the Raptor and the Raptor vacuum will be ready for full testing?

Online macpacheco

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #115 on: 03/21/2017 10:18 PM »
How far along is the Raptor engine?  Any word as to when the Raptor and the Raptor vacuum will be ready for full testing?
I'm no rocket scientist/engineer but it seems clear enough there will be a full year minimum testing before proper sea level / vacuum engines are produced for actual full thrust testing/qualification. The real for flight engines might not even be built in 2017.
This is still very early testing on a complete engine.
They will have to slowly increase thrust/change mixtures until the sub scale engine is running at its optimal (and more dangerous) parameters.
We don't know how much the engine components are finalized with margins to tolerate full power operations or a normal size engine.
I would wait at least until late summer/2017 to repeat such questions and hope for an actual answer.
Raptor is a crazy ambitious project. It not only intends to be one of the most efficient rocket engines in the world but also capable of 1000 mission firings (with at least 100 firings without any engine refurb). That and M1D are already good enough for current missions. They will take their time to do it right, much like M1C/M1D development progressed much slower than some people wanted, because Musk demanded the engine had crazy margins which are now paying off with Block IV/V thrust upgrades.
« Last Edit: 03/21/2017 10:20 PM by macpacheco »
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Offline dglow

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #116 on: 03/21/2017 11:31 PM »
A different angle: SpaceX isn't the only company building a methane SC engine. And they increasingly find themselves in direct competition, on multiple levels, with the other company doing so.

So not only do we not know, to any level of precision, the progress of Raptor development; I suspect we are unlikely to ever know much detail until the rocket is finished, or very nearly so. Blue is famously tight-lipped, and we've seen SX increasingly adopt a similar approach.

Sorry, that sucks as an answer. We can scout McGregor until the cows come home – or run away! – but we won't know Raptor is ready until either, a big Elon reveal (which won't necessarily coincide with 'finished'), or when we see it fly.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #117 on: 03/22/2017 01:08 AM »
Heck, SpaceX was more tight lipped than Blue Origin. Blue Origin did a press release with pictures and articles when the first BE-4 was finished, before even the first actual BE-4 test firing. SpaceX only showed the Raptor test firing. I think this may be because Blue has a customer that hasn't 100% decided on what engine to pick yet, so Blue has to make a big deal about any progress so it's obvious to all stakeholders. SpaceX just has themselves, in reality (other than some Air Force funding for development, which doesnt need public press releases).
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Offline Okie_Steve

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #118 on: 04/01/2017 11:09 PM »
Does anyone have a guestimate for the total wattage of a F9 S2 on orbit? It occurs to me that with Rapttor based restartable methalox upper stage engine as has been speculated, is might be worth while to include a methalox fuel cell to keep the batteries charged and/or  replace some of them for longer loiter time. Wondering how heavy it might have to be for the required power output compared to more/larger batteries

Offline Apollo100

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #119 on: 04/03/2017 10:54 PM »
Were the initial "Raptor" tests solely re-manufactured IPD hardware from AR drawings, or did they change the designs?

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