What's Up with Specific Impulse? 

Wow that was amazing Not as flashy as Tim Dodd’s video but even more informative and technical I managed to understand most of this surprisingly, thanks a lot for making this - I’ve binged watched 8 of your videos now and have to say your channel is very good

I googled specific impulse for dummies. This video is at the top of the list. Most helpful for me to understand the practicality (or lack thereof) of hydrogen as a rocket fuel. Hydrogen seems to be difficult to wrestle and corral in dummy speek. Thanks

And so ended the Teal Era of Eager Space!

Raptor Vacuum would be a better comparison against the RS-25, since it's also a more vacuum-optimized engine that can still function at sea level. It's Isp is ~300s at sea level and 378s in vacuum; the RS-25's corresponding values are both about 1.2x higher. The reasons the difference is smaller than the factor of 1.35 suggested by the difference in exhaust weight are twofold. First, as you stated, Raptor runs a significantly higher chamber pressure. Second, it also runs hotter; by about 200k.

Fantastic, spot on for those of us who... love the idea of rocketry.. but really don't have the brainpower for truly understanding it :D

Great explanation for such a complex topic! I wanna say something about 28:48 : How I see it, whilst your molecuar weight would decrease, at the same time your combustion chamber tempereature (and possibly pressure) would also decrease, since you would move even further away from the stochiometric mixture ratio. Combining these factors I think you would actually end up with lower Isp. (If you think about it, you are just reducing energy per unit mass of your propellant when going to lower Oxidizer to Fuel ratios here). Also, when comparing the RS25 with the nuclear thermal engine at 40:12, I think other factors could be the combustion chamber pressure or the expansion ratio, or maybe also looking at the combustion chamber temperature vs. just the exhaust temperature. But yeah I'm also not too sure how to explain that difference in Isp :)

Side-Note on Raptor running fuel rich: When watching a Starship launch, it is quite interesting to see that there is quite a bit smoke (or soot, I guess) coming off the exhaust. It becomes rather obvious from this, that not all the exhaust is either CO2 or H2O or indeed unburned methane, but some amalgamation of Carbons, too. Chemical reactions are rarely instantaneously perfect. I think, you don't quite see this effect as clearly with other common propellant types: pure Hydrolox cant have smoke therefore no visible impurities, kerolox obviously has lots smoke, solid fuels obvioulsy are pretty much nothing but smoke. Hypergolics, like hydrolox, tend to burn nicely into gasses, too, but I think you sometimes can see unburned particles (potentially from the pre-burner?), like on Proton or some of the Long Marches.

Thanks for this video. This is really awesome! There is one thing that I don't understand/ I am not satisfied with the given explanation about molecular mass directly correlating to Isp: In you comparison between the RS-25 and Raptor you explain that one reason why the RS-25 SL Isp is not much higher than Raptors, is because the RS-25 is optimized for use in vacuum and not for sea level. Of course, I agree with that overall, however, I think this is not the deciding factor in this discrepancy, because: a) the RS-25 over-expanded nozzle is rather uniquely designed as a sustainer engine that can avoid flow separation at sea level while maintaining its low gas exit pressure. Although I admit, I am not sure whether this benefits or worsens the Isp (flow separation is more of an issue for stability, not performance, I think). Theoretically, I would think lower exhaust pressure is even a benefit, no? b) I don't think there is any other conventional hydrolox engine with better SL Isp than the RS-25, even in concept. But I might be wrong on that. I found a Nasa document, stating the theoretical best Isp for Hydrolox would give 389.3s, which is still only 1,17 times better than Raptor, far from the expected 1.35. (Link to the document below, page 20, provided YT doesn't delete it again) c) also when we compare the RS-25 and Raptor Vac. version for their Vac. Isp we get 452.3s / 380s = 1,19. So it's again quite off the expected 1,35. Overall, your comparison/explanation doesn't add up. Or maybe I just don't understand it. To me it seems that hydrogen is just inherently not as good at sea level as it is in vacuum? But I am again not sure why that would be the case. Something about hydrogen exhaust velocity suffering from atmospheric pressure maybe? Or is it just that Raptor is a bad comparison, because of it's ridiculous chamber pressure? [Disclaimer: I am writing this comment before watching the section abut the RD-180 and everything after, so there is a chance that I'll eat my words in few minutes from now. Just wanted to get these thoughts off me before continuing :)] [Edit after finishing the video: I think it becomes very clear that molecular mass is very important but by far not the the sole factor for Isp. As correctly stated in the video, cycle types and ambient pressure optimization also play a big role. I think what explains the discrepancy between RS25 and Raptor best is indeed their difference in chamber pressure. If both ran at the same pressure, then we'd likely see results similar in difference to their respective exhaust gas masses. But with Raptor running at such high pressure it can overcome some of the disadvantages that methalox has when compared to hydrolox. You touch some of that in the section about RS-25 vs. SNRE, where RS-25 shows unexpectedly good results, probably due to its higher temperature and pressure. This just wasn't clear in the RS-25 vs. Raptor comparison. :) Or maybe I just got it wrong altogether?]

Love the video! One question I have is why not make all chemical propulsion systems with hydrogen if you can get such high ISP out of them? Supposedly you could get whatever thrust you wanted just by increasing mass flow, so why bother messing with RP-1 at all?

Absolut size of the nozzle of two different engines could be different, even if you compare engines with similar power , what matters is expansion ratio (that between nozzle neck hole and end of the bell), and density of the propellant or gasses exhaust, raptor and RS-25 work on different pressures(and with different propellants) , so isn't weird that that with higher chamber pressure is smaller. Other wise, I say is a good and detailed explanation of how those things work.

Mr. Rocket Sciense::: what minimum TWR, and acceleration, does a rocket need to launch just 200 meters up???

40:13 Nitpick: the degree symbol is used with Celsius, but never with Kelvin. I’m not sure why that rule was created. Anybody know?

17:44 SSME essentially had aerospike performance while having a bell nozzle

I believe that in case of the LH2 Nuclear thermal engine the H-H bond actually gets broken under the extreme heat so the reaction is H2 -> [Delta T] 2H and the weight of the exhaust is 1 g/mol

50:58 Everybody understands Isp?!? Nope, that's not true. We just get used to it.

How difficult can it be to count to 10? 18:10

47:00 NASA FAIRYTALES

earth's gravity = 9.8 m/s^2

The ‘no practicality’ of aerospike engines is often told but never proven, as a result of the cowardice of Aerojet Rocketdyne in refusing to continue designing them after the mid 90s. SpaceX may take advantage of its advantages with Starship due to the fact that Full flow engine cycle has some key similarities to another engine cycle often mentioned as benefiting the most from aerospikes: the Dual Expander. Both engine cycles use the bell as a contributor to heating the fuel. Both run their heated propellants through turbines before injection. But most importantly: both inject their circulating propellants as a gas (most of it in FFSC does pump out as liquid, unless I’m wrong.). I suspect future Raptor designs will employ altitude compensating nozzles as a means to reduce weight overall for all nozzles on the Ship (having a common engine instead of two altitude specialized variants should reduce weight and complexity, hypothetically, even if some weight is added to some nozzles and weight deducted from the others). It may very well make sense for lunar and Martian variants especially if they can be designed with smaller nozzles than a vacuum bell.

The exhaust velocity formula seems to be missing a factor of two in the sqrt? I looked at the wikipedia page of de laval nozzles for reference: en.wikipedia.org/wiki/De_Laval_nozzle