Quote from: baldusi on 11/08/2016 03:33 PMQuote from: rsdavis9 on 11/03/2016 06:42 PMI'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???

Quote from: rsdavis9 on 11/03/2016 06:42 PMI'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.

I'm super excited. But as you said no info to work with.

Are there any updates about Raptor development after the September test?

Elon Musk on Twitter:SpaceX propulsion just achieved first firing of the Raptor interplanetary transport enginehttps://twitter.com/elonmusk/status/780280440401764353Production Raptor goal is specific impulse of 382 seconds and thrust of 3 MN (~310 metric tons) at 300 barhttps://twitter.com/elonmusk/status/780275236922994688

Quote from: livingjw on 10/05/2016 12:39 PMQuote from: Dante80 on 10/04/2016 11:52 AMHere 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 358s4. 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).

Quote from: Dante80 on 10/04/2016 11:52 AMHere 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)

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?

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.

I don't know how to force a fixed width font in this forum.

Quote from: Manabu on 03/14/2017 06:17 PMI 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-OptimalExpansionERDiameter (m)Thrust (kN)Isp (s)Thrust (kN)Isp (s)H (km)P (atm)401.703037331.53274357.42.330.75340 V21.703052334.13287359.81.690.8152003.802315253.03500382.016.270.098200 V23.802361258.43536387.015.550.110Is that what you are after?

WARNING: Temperature T=93.00 K could not be assigned to the species "CH4(L)". Using T=298.15 K instead.

How far along is the Raptor engine? Any word as to when the Raptor and the Raptor vacuum will be ready for full testing?