Poll

When will full-scale hot-fire testing of Raptor begin?

Component tests - 2017
2 (1.1%)
Component tests - 2018
8 (4.4%)
Integrated tests -  2017
12 (6.6%)
Integrated tests -  2018
134 (73.6%)
Integrated tests -  2019
18 (9.9%)
Raptor is not physically scaled up
7 (3.8%)
Never
1 (0.5%)

Total Members Voted: 182


Author Topic: ITS Propulsion – The evolution of the SpaceX Raptor engine  (Read 148509 times)

Offline AncientU

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #660 on: 10/27/2017 01:18 PM »
...
But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.

Understand where you're going with this, but don't think Blue has the push-the-tech-to-the-limit DNA that is driving Raptor. 
Started with much lower goals and will end with much lower performance, IMO. 
Whatever you mean by 'effectiveness' (T/W, ISP, ?) will be lower, too -- lower than RD-180 and Raptor.
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Online Space Ghost 1962

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #661 on: 10/28/2017 01:53 AM »
...
But the effectiveness of the engines relative to RD-180 will be better, in every measure. Which means that likely Vulcan will also benefit downstream. Possibly NA as well.

Understand where you're going with this, but don't think Blue has the push-the-tech-to-the-limit DNA that is driving Raptor. 
Started with much lower goals and will end with much lower performance, IMO. 
Whatever you mean by 'effectiveness' (T/W, ISP, ?) will be lower, too -- lower than RD-180 and Raptor.

These teams all think they're doing the best engine, second to none.

But the first priority for BE-4 after reaching basic performance targets is ... reliable, reproducible, deterministic. Otherwise ULA can't use it. When they can use it, it doesn't have to be anything more than that to allow Vulcan to displace Atlas/Delta (with Centaur V ...). Doesn't have to eclipse Raptor/RD-180.

Yes, they could sit on their hands then. "Mission accomplished"  ::) But that's just the start for Vulcan/NG.

Given the processes they use, they retain the above, at no added risk, ... but increase chamber pressure/iSP/duration/margin. Vulcan vehicle/avionics/operations adapts to encompass this in missions.

RD-180 also has been gradually improving, as a mature design. But mature designs already have lesser bounds, as to go further, you have to add risk with significant changes, which likely exceed the scope of the business. (There is work on other RD-170 variants pressing.) What vehicles would those newer scoped engines fly in? Not Atlas/Vulcan.

Now circle back to Raptor - this thread.(BTW, all engine teams are acutely aware of the others work, practically in real time.) These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe. And the engines will have to on a single mission critically fire dozens of times without incident. Unlike BE-4/RD-180, who will have 2-3 critical burns in flight at max (Vulcan just one!). And since those burns will occur after extremely high delta-v targets (again unlike the other two), propulsion efficiency/iSP has to exceed the other two, to retain the advantage of the rocket equation with remaining propellant - to allow the vehicle architecture to realize its design goals.

Very different engines. BE-4, unlike RD-180 but like Merlin 1D, will have a vacuum variant. However, RaptorVac isn't in the same league. We're talking about optimization for in space propulsion as the majority of its role (the NG/NA architecture pursues hydrolox for this purpose), with a scaled engine to match that need without additional stages but with refueling.

(Note that ULA is backing off ACES and instead going for a expanded Centaur V. They can't get the "buy in" to fund the rest, which likely will be factored in incrementally as capability is desired.)

So they point to be made with this is that the different approaches by SX, BO, ULA in vehicles/engines is not in them being  more/less talented/aggressive/creative/experienced/... its instead the nature of what they are attempting to bring to bear.

For ULA its a next generation Atlas without past baggage, leveraging as much of the future as the parents will let them.

For BO its in entering the partially reusable LV business at Ariane/SX level of capability/flight frequency.

For SX its in a fully reusable vehicle with interplanetary HSF capability in excess of SLS.

Back to RD-170 variants for comparison - likely enhanced single chamber for Angara, methalox four chamber derivative for Russian SHLV like SLS. They'll not need much more than per chamber what RD-180 already does. And reuse isn't yet on the map.

Offline rsdavis9

I think I know the answer.
Quote
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
With ELV best efficiency was the paradigm. The new paradigm is reusable, good enough, and commonality of design.
Same engines. Design once. Same vehicle. Design once. Reusable. Build once.

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #663 on: 10/28/2017 12:44 PM »
I think I know the answer.
Quote
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
Full flow means you are extracting more energy for pumping.
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Online Space Ghost 1962

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #664 on: 10/28/2017 11:17 PM »
I think I know the answer.
Quote
These guys are after extreme chamber pressures ORSC of any kind ... could never reach. Energy densities that are not on the development plan anywhere else on the globe.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).
Full flow means you are extracting more energy for pumping.
Advantages of a full-flow staged combustion cycle engine system

The chief point here is that the separate OR/FR paths are at less pressure than the combustion chamber, and there is no interpropellant seal to fail at extreme pressure or transient flow. The maximum chamber pressure is thus set by the design limits of the combustion chamber and injector(s).
« Last Edit: 10/28/2017 11:22 PM by Space Ghost 1962 »

Offline ChaoticFlounder

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #665 on: 10/28/2017 11:38 PM »
I think I know the answer.
Why is a FFSC higher chamber pressure than a OSRC?
I assume it has something to do with 2 gas flows(FFSC) going into the chamber versus one liquid and one gaseous(OSRC).

Hopefully this will help you understand a little bit better the design choices they're making

so, pump power requirement is directly proportional to mass flowrtate and output pressure (P1)

P1 is going to be your highest pressure in the system and will directly dictate your chamber pressure (Pc)

so, turbine work per unit mass is equal to

Tw/m = Cp*(T4 - T5)

where:

Cp = Specific Heat
T4 = Turbine Inlet Temperature
T5 = Turbine Outlet Temperature

This is pulled directly from this site:  https://www.grc.nasa.gov/www/k-12/airplane/powtrbth.html

When you multiply this specific turbine work by your preburner mass flow rate, you get shaft power (Ps)

This shaft power is used to drive the pump;  therefore, pump output pressure is directly related to (preburner mass flow)*(deltaT) across your turbine

deltaT is limited by your materials, so your variable to change is mass flow and that is what the FFSC does in spades

Essentially, it's saying that your theoretical pressure limit is going to be around twice what your theoretical pressure limit will be with FRSC / ORSC <- assuming T4 is the same and your propellant densities are similar

All, please correct me if I said anything wrong and ask if you have questions.

C



Offline IainMcClatchie

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #666 on: 11/11/2017 08:03 AM »
How hot is the partially combusted oxygen in the turbopump?

In the following quick analysis, I'm assuming a chamber pressure of 25 MPa and that the turbine and compressor are 100% efficient, with no heat leakage into the incoming LOX at 66 K.  I tried a range of preburner pressures from 35 to 60 MPa.  I got very, very low temperatures.

The output of the compressor is a subcritical liquid around 72 K.  The preburner appears to heat it up to at most room temperature.  At 50 MPa turbine inlet temperature, the turbine inlet is at -17 C.  Flameholding will be a challenge.  The preburner is going to need to burn with a small fraction of the LOX before mixing the result with the bulk of the LOX, a more radical version of the burner cans in turbofan engines.

There is a substantial benefit to running the preburner at high pressure: the turbine extracts more energy, and so less propellant is burned in the preburner.  That means everything from the preburner output to the injector face gets more dense and therefore smaller.  In particular, the volume gets smaller faster than the pressure goes up, so the figure of merit for a pressure vessel, which is pressure*volume, goes DOWN at higher preburner pressure, while at the same time you get a small bump in exhaust velocity from burning more of the propellant in the main chamber..

That means an engine with a higher preburner pressure can be lighter weight, which seems counterintuitive to me.

Turbine energy extracted varies from 7% at 35 MPa to 20% at 60 MPa.

I wonder how useful it is to fully mix the preburner.  More energy can be extracted from a hotter stream, so that for a given chamber pressure, less propellant can be used by the preburners and a higher chamber temperature and better exhaust velocity can be achieved.  With these ridiculously low temperatures, there is gobs of temperature headroom.  It seems totally feasible to triple the energy extracted by the turbopump.

I don't think they are going to need ceramic coatings for this thing.  The turbine bits do have to deal with high pressure oxygen, but surely that's more benign than hot oxygen.

I've attached my spreadsheet if anyone wants to check my math.

Offline John Alan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #667 on: 11/11/2017 05:19 PM »
Looking at the above... I'm kind of surprised and then again not...  ???

I mean, my take on this FFSC cycle is... Use the "energy" stored in the liquefied prop as it "boils" back to a vapor to drive a turbine powering the pump...

I always figured the highest pressure (2x+ chamber) would be found between the pump outlet and the "heater" in line...
I don't like to call it a preburner... as we are not sure it actually will use much if any of the LOX flow rushing by...
I like to simplify it, by picturing a burner can fed gaseous oxygen and gaseous methane from other onboard sources...
The vaporizers for these flows MAY be the burner can itself separately fed high psi liquid prop from upstream (so it can be controlled)...
The exhaust from the preburner mixes with the cold LOX and heats it enough to flash it all to oxygen "steam"...
It's still at nearly the same high pressure as it enters the turbines and the pressure is then dropped across the turbines converting the energy of the much expanded flow into mechanical energy to drive the pumps... 

LOL... Yes... it's kind of like taking a fire hose and hooking it up to a steam turbine...   ;D
Just got to put a big enough heater in line to convert all the water to steam before it reaches the turbine...  :o

Anyway... that's how I wrapped my head around FFSC Raptor inner workings...  ;)
« Last Edit: 11/11/2017 05:25 PM by John Alan »

Offline Robotbeat

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #668 on: 11/11/2017 06:10 PM »
That doesn't make sense.
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Offline yokem55

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #669 on: 11/11/2017 07:21 PM »
SpaceX Veterans Day Commemoration pic on Twitter has the McGregor vets pictured in the Raptor Test cell.


https://twitter.com/SpaceX/status/929441494494208000
« Last Edit: 11/11/2017 08:27 PM by gongora »

Offline IainMcClatchie

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #670 on: 11/11/2017 07:52 PM »
It seems the preburner has to mix the flame exhaust into the LOX pretty well to avoid two-phase hitting the turbine.  So much for making the turbine more efficient.

That said, the oxygen preburner is using 2-3% of the total energy flow of the propellant, most of that just for the phase change rather than temperature change.

Note that the oxygen is hardly *flashing* to "steam".  This diagram shows that the transition from liquid to supercritical is very smooth at 25 MPa.  50 MPa will be very smooth as well.

Neither oxygen nor methane is squishy in their liquid forms, so the turbopump shaft power is 15-20 megawatts instead of some even more ludicrous number necessary for a hydrogen compressor.

The turbopumps in the Raptor will be amazingly small.  The fluid density coming from the preburner will be well over half that of water, so the volume is under 800 liters/sec.  I'm not sure what the right velocities are, but at 100 m/s, that's a cross section of 80 cm^2.  I doubt it ever goes in a pipe but if it did it would be 10 cm diameter.

Does anyone have some rules of thumb for determining the turbopump compressor and impeller sizes?  The tip speed needed for a single-stage 60 MPa centrifugal pump is 304 m/s, which is just below the speed of sound in cold LOX.  Liquid methane has a much higher speed of sound, well over 1 km/s.  If a single-stage impeller can do the job, I'm wondering if the turbopumps can eliminate the shaft and stator vanes entirely and just consist of an impeller and turbine back-to-back, with the burner cans around the periphery.

Offline John Alan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #671 on: 11/11/2017 08:13 PM »
Agreed... "Steam" was a term used in gross error...
BUT, I was about to use a water analog, so I used it to help connect the two...  :-[

I understand gas turbines and steam turbines (air and water working fluids) and was just trying to relate it to a LOX working fluid...

My bad...  :P

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The aforementioned test stand pic.
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Offline docmordrid

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #673 on: 11/12/2017 04:16 AM »
The aforementioned test stand pic.

Full-scale engine?
DM

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The aforementioned test stand pic.

Full-scale engine?

It's hard to judge from this picture, and the difference between subscale and full scale is going to be extremely small, ~10-20% as modelers on this thread have estimated. I expect Musk will tweet about it whenever SpaceX conducts the first "full scale" tests or at least the first 250 bar test.

But he suggested that the main hurdles between the test article and flight engine would be aggressive mass reduction and reliability improvements, so no guarantee that anything more than the test firing videos we've been given will be deemed tweetworthy.
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Offline Hominans Kosmos

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #675 on: 11/12/2017 01:03 PM »
Re: cold turbine exhaust temperatures.

It appears to me you want the turbine exhaust at all thrust regimes to be above the condensation temperature of steam. At 20..30 bar steam should condense below 230..240 celsius.

After the preburner/heater the propellant is a mix ture including combustion products.

No?
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Offline philw1776

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #676 on: 11/12/2017 02:32 PM »
The aforementioned test stand pic.

Full-scale engine?

L2 discussion on that very issue
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Offline John Alan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #677 on: 11/12/2017 05:35 PM »
Re: cold turbine exhaust temperatures.

It appears to me you want the turbine exhaust at all thrust regimes to be above the condensation temperature of steam. At 20..30 bar steam should condense below 230..240 celsius.

After the preburner/heater the propellant is a mix ture including combustion products.

No?

Yes and No... (my opinion)
Yes there will be some small amount of water vapor in the combustion products downstream of the burner can...
BUT...
Note the heat added to the stream is only enough to phase change the easily pumped, not compressible, sub cooled LOX into squishy. expandable, but supercritical at high pressure gaseous Oxygen  in good enough condition to be expanded over a turbine section from say 600 bar down to 350 bar on it's way across say a final 50 bar drop across the chamber injector into a running, firing 300 bar rocket combustion chamber...
Literally... the temps at the oxygen turbine may be room temperature... and the turbine design will have to allow for some liquid droplet (lox or otherwise) to pass thru harmlessly I believe...

The startup sequence on this must be very interesting... I must say...  ???

The fact the above picture of a pristine Raptor development test stand with a year old development engine  (assuming same basic assy) still in place is a HUGE achievement...
Tom Mueller and his group figured out how to start and stop this thing without it blowing up...  8)

Later edit... fixed my bar numbers above off by factor 10 (woops)
« Last Edit: 11/12/2017 11:23 PM by John Alan »

Offline Hominans Kosmos

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #678 on: 11/12/2017 05:46 PM »
From my understanding water droplets in a gas turbine is a bad thing, abrasive blasting or some such effect.
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Offline John Alan

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Re: ITS Propulsion – The evolution of the SpaceX Raptor engine
« Reply #679 on: 11/12/2017 05:50 PM »
From my understanding water droplets in a gas turbine is a bad thing, abrasive blasting or some such effect.

Yes it is... very much so in both gas and steam turbines with their sharp and delicate edges that are easily damaged by water or ice...
Somehow they must have a design that can handle it...  ;)

On edit...
I am also quite sure this design and control information will never see the light of day as long as SpaceX has say over their designs and intellectual property... BUT, we can speculate...  :)

Much later edit...
I'm wondering if maybe the turbine has the reverse of gas turbine industry typical and has passageways in the blades to allow HEATING the turbine somewhat using hot gas tapped from the upstream burner assy so as to heat the blades enough to vaporize on contact any stray droplets that find their way into contact with the turbine...  ???
If so... that's some interesting stuff right there...  ;)
« Last Edit: 11/12/2017 07:25 PM by John Alan »

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