New Mars Forums

Another article (long) on recent work on making the warp drive feasible:

MAY 27, 2022 BY MATT WILLIAMS
The Dream of Faster-than-Light (FTL) Travel: Dr. Harold “Sonny” White and Limitless Space.
https://www.universetoday.com/155995/th … ess-space/

  Robert Clark

I had been investigating other topics unrelated to space. But the announcement of the Radian Aerospace SSTO venture piqued my interest. I'll start a separate thread on it discussing the SSTO aspects.

  Robert Clark

What do you mean "we"? 

  Bob Clark

Thanks for the response. Going the high performance engine and lightweight tanks route can double the payload. Note that SpaceX spent quite a bit of time and money to double the F9 payload with the F9 Full Thrust. This would be an alternative way of accomplishing the same thing. Quite key to remember is that a high performance engine can be accomplished by using an existing midlevel performance engine such as the RS-68 on the Delta IV or the Merlin on the Falcon 9 by adding an altitude compensation nozzles. You don't need an expensive engine such as the SSME.

Here's one simple way, low cost way of getting altitude compensation: flexible, extensible, nozzle attachments:

So launch companies can double the payloads of their TSTO by taking this approach. BUT additionally it would also allow their first stage alone to have significant payload to LEO. Note most payloads don't need the full launchers payload capacity to space. Then with the SSTO you could launch the smaller payloads more cheaply without needing the upper stage at all.

  Robert Clark

Thanks for the interesting link.

  Bob Clark

Here’s a video clip of the Raptor Vacuum being tested on the ground:

https://twitter.com/erdayastronaut/stat … 26240?s=21

At a 107 to 1 expansion ratio, it has a shorter nozzle than the expected operational version. Someone in that twitter discussion suggested it had a support ring placed around the nozzle to deal with the high side loads.

  Bob Clark

Thanks for that. For the reusable rocket application, the “no-moving-parts throttle“ looks to be the way to go to avoid the extreme heating issues in the throat for a rocket engine.

SpaceX also wants high throttle ability for landing, so perhaps you could provide them advice there:

https://twitter.com/search?q=Throttle%2 … ery&f=live

  Bob Clark

GW, I’m interested in any and all methods of altitude compensation, from the aerospike, to methods of physically changing the nozzle size, to any other method.

One interesting possibility is to not vary the nozzle size at all but to vary the throat size, keeping in mind it’s the nozzle area to throat area ratio that needs to vary.

So I was interested to read this comment of yours:

Advanced Aerospike Rocket Engine Design
These tactical nozzle designs included both straight conical and full curved bell designs.  We never tried anything else as a production item.  The flapper-lollipop throat insert as a variable-geometry nozzle for ASALM is the exception.  It worked,  but the real-world design restrictions were twofold:  (1) the wake zone behind the lollipop had to close before the exit plane if full nozzle efficiency was to be obtained,  and (2) you could not put a large pressure drop across a structure like that because there were (and are) no materials capable of withstanding such abuse.
http://newmars.com/forums/viewtopic.php … 80#p146380

Could you expand on the problems you observed with that?

   Bob Clark

I got the spread sheet. Thanks. It’s very useful for doing estimates. You might want to make it available on Google Docs so anyone interested in doing the calculations of engine performance with various chamber pressures and expansion ratios can do them.

I have a question about the example “Conventional 3”. It discusses a fixed nozzle, I assume, kerolox engine at a 1800 psi, 120 bar, chamber pressure and 35.5 nozzle area ratio. This results in a surprising high 341 s vacuum Isp. This would mean a sea level fireable engine reaching the high vacuum Isp of the Merlin Vacuum upper stage engine.

However, the spreadsheet notes the flow separation would be close to happening at sea level. Likely though a slightly higher chamber pressure would make this less likely to happen. This is interesting because I did a calculation using GW’s criterion in post #28 for when flow separation could occur applied to the high performance Russian kerosene engines the RD-180 and RD-191, and was surprised to see their chamber pressures were so high at ca. 260 bar, that you could put larger nozzles on them to get 360s vacuum Isp yet would still be fireable from sea level, i.e., no flow separation.

If this is correct then these engines all you would have to do is put longer fixed nozzles on them and you would have SSTO capable engines, no altitude compensation required. This is surprising if true. Because the higher vacuum Isp would be important to get higher payload for the standard TSTO as well. So why didn’t the Russians give the engines the larger fixed nozzles to begin with? 

This example “Conventional 3” does have a somewhat higher chamber pressure at 120 bar compared to the Merlin’s 100 bar, though. It would be interesting to do the calculation to see how much payload the “Falcon 9” could get to orbit with this engine instead of the Merlin’s. Note also with this fixed nozzle it could be SSTO with significant payload, no altitude compensation required.

I put “Falcon 9” in quotes because this is a much larger nozzle size than the Merlin so 9 wouldn’t fit underneath the same diameter booster. So you would have to design engines with higher thrust so a fewer number of them could lift the same size vehicle.

Also, note that looking at example “Conventional 1” with the smaller nozzle area ratio 13.8 to 1. This is close to the Merlin 1D sea level engine’s nozzle area ratio of 16 to 1. But its sea level Isp is much higher at 303 s compared to 282 s. And it’s vacuum Isp is also better than the Merlin 1D at 324 s compared to 311 s.

This illustrates the usefulness of alt comp. It could improve the sea level Isp to 303 s by using the sea level area of 13.8 to 1 while giving the 341 s vacuum Isp by having the adaptive nozzle expand to 35.5 to 1 in vacuum.

   Bob Clark

This is actually the crux of the matter. Even a 10% increase in Isp can result in a 100% increase in payload for a SSTO. That is the nature of exponential change. Most people interested in space have heard the phrase, “the tyranny of the rocket equation.”
The other side of the coin is a small change in Isp can result in a radical change in payload. I like to call this “the beneficence of the rocket equation.”

  Bob Clark

Thanks for the calculation, but quite key is even for the first stage a significant portion of the flight is in near vacuum conditions. That is why first stage engines are overexpanded, with a nozzle size larger than optimal for sea level operation to be able to get better performance in vacuum.

For instance, the F9 first stage engine cuts off after about 2 and 1/2 minutes at about 60 km altitude. At that altitude and even before a vacuum optimized engine would be getting close to full vacuum Isp. So most of the 9 minute or so flight to orbit is actually at near vacuum conditions, so would be getting close to full vacuum Isp of a vacuum optimized engine.

Then the average Isp of an alt. comp. kerolox engine would be close to the 360 s max Isp of a vacuum optimized engine. Actually, the max Isp could probably even be in the 370s to 390 s range. But 360 s is in the range achieved for upper stage kerolox engines:

http://www.friends-partners.org/mwade/p … xosene.htm

Likewise, for a methanolox alt. comp. engine, the average Isp would be close to the 380 s max Isp of the vacuum optimized engine.

Here’s an image illuminating most of the flight to orbit place where the Isp would be near the max Isp level:

This is for hydrolox, though.  I’ve not seen any analysis of a dense fuel engine using alt. comp.

  Bob Clark