CNES ESA Prometheus / Callisto proposal

[FutureSpaceTourist]

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Europe, Japan Plan 2021 Reusable Launcher Demo
Aerospace Daily & Defense Report
Thierry Dubois
Oct 25, 2018

LYON, France—French space agency CNES has released details on the reusable launcher demonstrator it plans to test from 2021 in Kourou, French Guiana, with its German (DLR) and Japan Aerospace Exploration Agency (JAXA) counterparts.

http://www.aviationweek.com/space/europe-japan-plan-2021-reusable-launcher-demo

[FutureSpaceTourist]

Prometheus: Demonstrator of Future Engine passed its Definition Review
Paris, 4 February 2019 

* ArianeGroup has just finalized the Definition Review of the, of the Prometheus engine demonstrator, on 1 February 2019 with the support of European Space Agency, CNES and DLR
* It demonstrates the pertinence of the design and the technological choices made and confirms the program’s ambitious cost objectives
* Prometheus is a European demonstrator for a very low cost and potentially reusable engine
* The bench tests of the first two examples of the engine are scheduled for as early as 2020

[...]

https://www.ariane.group/en/news/prometheus-demonstrator-of-future-engine-passed-its-definition-review/

Full press release attached.

[FutureSpaceTourist]

https://twitter.com/arianegroup/status/1096458141699661824

Prometheus is a precursor of the future engines intended for use by Europe’s launchers by 2030! 👇🏻

[Salo]

https://elib.dlr.de/120200/1/SP2018_478_PA.pdf

3.1.2 LOX-LCH4 engines

Several initiatives are currently working on engines with the propellant combination LOX-Methane. Although proposed several times in the past, this “softly cryogenic” blend has never yet been realized in an operational launcher stage.
The main combustion chamber MRs of this combination have been selected close to their optimum Isp, however, slightly shifted towards increased MR to reach increased bulk density. This approach is different to the LOX-LH2 engines and results in slight differences in MCC-MR and significant differences in the engine MR. The method is used in a similar way also for the other hydrocarbons and is justified by their increased propellant density but considerably lower mass specific impulse compared to LOX-LH2.
The gas generator operates methane-rich and its hot gas powers the single shaft turbine. Major characteristics are derived of the PROMETHEUS-Demonstrator [7] but the baseline assumptions remain similar to all other engines of the system study. Obtained data (Table 2) are not far off the expected PROMETHEUS-engine.
The staged combustion type is based on a fuel rich preburner design with a single-shaft turbopump. It’s worth noting that both simulation tools lrp and RPA converged only for relatively high preburner pressures resulting in lower T/W than other engines. A direct comparison with another engine is not possible because the staged combustion methane engines under development in the US, Raptor and BE-4, intend to operate in FFSC and in LOX-rich-mode and at significantly different chamber pressures [6]. The LOX-Methane engines deliver the highest performance of all hydrocarbon types, yet roughly 80 s to 90 s below the LOX-LH2 engines.

[Salo]

Earlier on this article authors was more optimistical:
https://elib.dlr.de/114430/1/Paper_IAC2017_D.2.4.3ENTRAIN.pdf

[envy887]

Something odd is going on in these studies. They are designing vehicles larger than New Glenn, doing the same downrange landing, but getting less than half the expected payload (even compared to the 2-stage methalox NG). And they are nearly 3 times the GLOM of F9 with only 30% more payload.

[ZachF]

Something odd is going on in these studies. They are designing vehicles larger than New Glenn, doing the same downrange landing, but getting less than half the expected payload (even compared to the 2-stage methalox NG). And they are nearly 3 times the GLOM of F9 with only 30% more payload.

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

A Methane Merlin should be easily able to get an ISP of 300 sl, of course, methane doesn't really begin to shine until you use staged combustion and densified propellants.

I have modeled various architectures for first stage recovery, and my experience has been that once average ISP drops below ~300 (sl&v) stage recovery enters a pretty steep region of diminishing returns.

This study is looking more and more like it was goalseeked; where data is fit to an outcome they already pre-determined.

[ZachF]

@Envy887

Here is why their proposal performs so bad. I swapped out Starship's superheavy with a Prometheus-powered lower stage with the same performance requirements.

Basically, it's what would be required to push 1,354 tonnes to a delta V of 3,457m/s(without losses), then it needs to perform a dV of ~3,348m/s to itself to land.

Using non-desified propellants and a mixture ratio of 2.5:1 vs ~3.7:1 reduces fuel density from ~0.9 to 0.777.

Lower TWR of engines plus lower density fuel likely reduces unfuelled mass from ~4.5% on starship to >6% for Prometheus stage.

In short, first stage mass increases by a whooping 104%!

First stage fuel volume increases by a whopping 131%!
 
Empty mass of first stage is almost tripled!

With these stats an expendable Prometheus stage is probably worse than a reusable Raptor stage!

[envy887]

Something odd is going on in these studies. They are designing vehicles larger than New Glenn, doing the same downrange landing, but getting less than half the expected payload (even compared to the 2-stage methalox NG). And they are nearly 3 times the GLOM of F9 with only 30% more payload.

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

A Methane Merlin should be easily able to get an ISP of 300 sl, of course, methane doesn't really begin to shine until you use staged combustion and densified propellants.

I have modeled various architectures for first stage recovery, and my experience has been that once average ISP drops below ~300 (sl&v) stage recovery enters a pretty steep region of diminishing returns.

This study is looking more and more like it was goalseeked; where data is fit to an outcome they already pre-determined.

The mass fractions are about 30% worse than Falcon. This is about right without densified propellant, but I don't know why they wouldn't baseline densified propellants. They aren't going to get very far without pushing some limits.

[ZachF]

Done the math... with those stats, an expendable Prometheus stage is worse than a reusable Raptor stage!

Threw in a BE-4 powered stage for comparison. Raptor is again a large improvement here too.

[ZachF]

Something odd is going on in these studies. They are designing vehicles larger than New Glenn, doing the same downrange landing, but getting less than half the expected payload (even compared to the 2-stage methalox NG). And they are nearly 3 times the GLOM of F9 with only 30% more payload.

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

A Methane Merlin should be easily able to get an ISP of 300 sl, of course, methane doesn't really begin to shine until you use staged combustion and densified propellants.

I have modeled various architectures for first stage recovery, and my experience has been that once average ISP drops below ~300 (sl&v) stage recovery enters a pretty steep region of diminishing returns.

This study is looking more and more like it was goalseeked; where data is fit to an outcome they already pre-determined.

The mass fractions are about 30% worse than Falcon. This is about right without densified propellant, but I don't know why they wouldn't baseline densified propellants. They aren't going to get very far without pushing some limits.

Poor mass fractions and poor ISP = crap system for reusability.

Ariane, unfortunately, is just looking for the easy way out.

The more you do the math, the more you realize Raptor's combination of densified propellant, good ISP, and good TWR all come together to make a system that make reusability much easier.

[Oli]

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

Merlin 1D has a sea level ISP of more than 289s?

[ZachF]

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

Merlin 1D has a sea level ISP of more than 289s?

I guess it's just equal (for some reason memory thought Merlin 1D was 292...)

SL ISP of 289 plus fuel density of 0.77 is pretty objectively terrible for a first stage engine though. They are going to have a hard time with reuse with those numbers.

[Oli]

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

Merlin 1D has a sea level ISP of more than 289s?

I guess it's just equal (for some reason memory thought Merlin 1D was 292...)

SL ISP of 289 plus fuel density of 0.77 is pretty objectively terrible for a first stage engine though. They are going to have a hard time with reuse with those numbers.

Here

https://spacelaunchreport.com/falcon9ft.html

it says 283s/312s, so 289s/320s for a methalox GG engine isn't terrible, given that Merlin is a highly optimized engine that went through several iterations.

IMO the study if flawed because it doesn't allow for a LH2 second or third stage for high energy missions. LEO would be the primary market for such a launcher.

[ZachF]

Look at the "key performance data" posted above by Salo... They are only modelling a sea level ISP of 289 for their methane engines(!) worse than the current kerolox Merlin... and a mixture ratio of 2.5 and TWR under 100. Of course that rocket is going to suck!! 

Merlin 1D has a sea level ISP of more than 289s?

I guess it's just equal (for some reason memory thought Merlin 1D was 292...)

SL ISP of 289 plus fuel density of 0.77 is pretty objectively terrible for a first stage engine though. They are going to have a hard time with reuse with those numbers.

Here

https://spacelaunchreport.com/falcon9ft.html

it says 283s/312s, so 289s/320s for a methalox GG engine isn't terrible, given that Merlin is a highly optimized engine that went through several iterations.

IMO the study if flawed because it doesn't allow for a LH2 second or third stage for high energy missions. LEO would be the primary market for such a launcher.

That's for 1D, 1D Full Thrust opened the throat a little to increase thrust. That lowered the ER, which slightly increased SL ISP while slightly decreasing vac ISP. Chamber pressure also increased slightly.

http://www.b14643.de/Spacerockets/Specials/U.S._Rocket_engines/engines.htm

1D FT is 289/312. Average ISP for both didn't really change much though.

Methalox just doesn't work well without staged combustion and densification.

[Oli]

That's for 1D, 1D Full Thrust opened the throat a little to increase thrust. That lowered the ER, which slightly increased SL ISP while slightly decreasing vac ISP. Chamber pressure also increased slightly.

http://www.b14643.de/Spacerockets/Specials/U.S._Rocket_engines/engines.htm

1D FT is 289/312. Average ISP for both didn't really change much though.

Methalox just doesn't work well without staged combustion and densification.

Interesting, thanks. In that case ~303s/334s would make more sense, especially since they assume a chamber pressure of 12MPa, which is more than Merlin. Weird.

[ncb1397]

Something odd is going on in these studies. They are designing vehicles larger than New Glenn, doing the same downrange landing, but getting less than half the expected payload (even compared to the 2-stage methalox NG). And they are nearly 3 times the GLOM of F9 with only 30% more payload.

Where did you get 3 times the GLOM of F9? Anyways, New Glenn has a gigantic expendable upper stage. And the diameter and length are smaller than New Glenn (5.5-5.2 m diameter, similar length).

edit: It appears that you were looking at the model based on structural indices that over-estimates 2nd stage mass(and therefore 1st stage mass).

edit 2: And propane?

[ZachF]

That's for 1D, 1D Full Thrust opened the throat a little to increase thrust. That lowered the ER, which slightly increased SL ISP while slightly decreasing vac ISP. Chamber pressure also increased slightly.

http://www.b14643.de/Spacerockets/Specials/U.S._Rocket_engines/engines.htm

1D FT is 289/312. Average ISP for both didn't really change much though.

Methalox just doesn't work well without staged combustion and densification.

Interesting, thanks. In that case ~303s/334s would make more sense, especially since they assume a chamber pressure of 12MPa, which is more than Merlin. Weird.

A GG methalox engine that gets 289/320 is probably around 70 MPa bar in chamber pressure.

EDIT: fixed unit mistake

[envy887]

Something odd is going on in these studies. They are designing vehicles larger than New Glenn, doing the same downrange landing, but getting less than half the expected payload (even compared to the 2-stage methalox NG). And they are nearly 3 times the GLOM of F9 with only 30% more payload.

Where did you get 3 times the GLOM of F9? Anyways, New Glenn has a gigantic expendable upper stage. And the diameter and length are smaller than New Glenn (5.5-5.2 m diameter, similar length).

edit: It appears that you were looking at the model based on structural indices that over-estimates 2nd stage mass(and therefore 1st stage mass).

edit 2: And propane?

Figure 7 of the study posted below. F9 has a GLOM of about 550 t. The methalox and propalox vehicles have a GLOM of ~1400-1750 t.

Even New Glenn is going to be in the 1400 t range for GLOM, since it will only have ~1700 t of thrust at liftoff. It does have the advantage of staged combustion, but it is using a first-iteration engine from a private company that never developed a staged combustion engine before. Surely ESA and DLR with all of Europe's resources could develop something similar.

[envy887]

That's for 1D, 1D Full Thrust opened the throat a little to increase thrust. That lowered the ER, which slightly increased SL ISP while slightly decreasing vac ISP. Chamber pressure also increased slightly.

http://www.b14643.de/Spacerockets/Specials/U.S._Rocket_engines/engines.htm

1D FT is 289/312. Average ISP for both didn't really change much though.

Methalox just doesn't work well without staged combustion and densification.

Interesting, thanks. In that case ~303s/334s would make more sense, especially since they assume a chamber pressure of 12MPa, which is more than Merlin. Weird.

A GG methalox engine that gets 289/320 is probably around 70 MPa in chamber pressure.

You mean 70 bar, or 7 MPa? That is what I get in RPA lite for a 70 bar methalox engine at 3.25 O/F, assuming 0.966 engine efficiency and 3% mass flow through the GG.

Using the same assumptions with 12 MPa,  I get 301 SL / 332 vacuum.

To get 320 seconds in vacuum requires using 6.7% of the mass flow in the GG, which seems really high.