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Separation-limited vacuum engines (like the current vacuum Raptor) inherently have utterly-lousy sea level thrust!  There is simply no way around that!  3x275 + 3x175 = about 1350 tons with all 6 burning at sea level on "Starship".  That's thrust/weight only 1.02 at liftoff,  which is long known to correspond to gravity losses WAY TO HELL-AND-GONE ABOVE 20% (or more) of LEO speed,  not the 5% of an efficient system.  Add only 30 tons of payload to this example,  and this thing CANNOT budge a single inch off the launch pad,  no matter how much propellant it has! 

And there is NO ROOM behind it for more engines!  Making the tankage hold 1300 or even 1400 tons really does not change that picture very much at all.

All SSTO designs face exactly the same thrust problem as trying to make an SSTO out of "Starship"!  You cannot have any more engines,  because those added would lie outside the stage diameter!  That doubles-or-more your drag,  and way-more-than-doubles your drag loss,  which with a really clean shape of the right L/D ratio is about 5% of LEO speed. 

There is simply way-far-more to this entire question than just Isp and mass ratio in the rocket equation!  I have long tried to communicate that,  but unsuccessfully!

And by the way,  if sea level thrust gets reduced by the backpressure term,  so does the corresponding sea level Isp,  for the same combustion chamber design and total propellant flow rate.  Which is EXACTLY why you need to look at engine/nozzle ballistics,  and not just pull Isp's out of some table in some reference.

I have provided the spreadsheet tools and the instructional lessons,  for free,  to be able to do this work correctly.  That's the stuff accessed by links posted right here on these forums. 

GW

Design by smashing may not always work as SpaceX May Be Failing to Get Starship Working at All
"All flight 9 has proven is just how much of a dead end Starship is."

SpaceX is still a long way from achieving reusability or a hefty payload capacity. Regarding the former, last week's flight test was the first time SpaceX reused a Super Heavy booster, replacing four of its 33 engines. Regarding the latter, the test had Starship carry a dummy payload of a measly 16 metric tons. Musk has previously promised that Starship will carry 150 tons.

Maybe its time to step back and try for the mid sized rocket....

tahanson43206,
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Inputs
Launch Vehicle: User Defined
Number of Stages: 1
Strap-on Boosters?: No
Dry Mass: 26,372kg
Propellant Mass: 2,185,289kg
Thrust: 33,854kN
Isp: 304.2s (90% of Vacuum Isp for the RD-180)
Default Propellant Residuals?: Yes
Restartable Upper Stage?: No
Payload Fairing Mass: 0kg
Launch Site: Cape Canaveral (USA)
Destination: Earth Orbit, Apogee 185km, Perigee 185km, Inclination: 45 degrees

Outputs
Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 45 deg
Estimated Payload: 78,854kg
95% Confidence Interval: 58,892kg - 103,231kg,922
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LOX/RP1 to Deliver 6,508,946,390N-s
Real World Engine Proxy: RD-180 engine's Isp (90% of Vacuum Isp) / thrust / mass flow performance figures?:
LOX: 1,597,846kg (1,192.422m^3)
RP1: 587,443kg (683.074m^3)
Total: 2,185,289kg (1,875.496m^3)

2,301,409kg LOX/RP1 + 116,120kg (vehicle and useful payload) = 2,301,409kg GLOW
116,120kg = 5.05% of GLOW
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For RGClark ... I get the impression you only read a tiny bit of the forum, and posts intended for you never reach you.

There is nothing wrong with that. That is how a lot of our members use the forum.  However, what it means is that feedback someone might offer is never seen, so if correction is offered you continue on without it.

(th)

By my estimates, the mass ratio of a LOX/RP-1 SSTO would be about 15.  Which means that the entire vehicle and payload would need to be no more than 6.7% of takeoff mass.  That is a very thin structural margin, especially if it includes thermal reentry shielding.  Some kind of launch assist would appear to me to be a necessity for this concept to have a chance of working.  If we can shave 30% off of the fuel required to reach orbit, then a reusable LOX/RP-1 SSTO can get to orbit with a more achievable mass ratio of about 11.  That is still tough to do.  But it is possible.

Another option would be a two staged vehicle, with both stages being reusable.  This is exactly what SpaceX are doing.

Dr Clark,
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Adding or subtracting 10 seconds of Isp is almost meaningless for a LOX/LH2 engine.  A straight Delta-V calculation will show that the difference between 452.3s of Isp and 464.9s of Isp is 85.65m/s of additional acceleration possible for the same propellant load and dry vehicle mass.  You will not see any sort of night-and-day payload performance improvement by using that extendible nozzle with the RS-25, but your vehicle will have significant additional dry mass.  Unless some other part of your vehicle becomes lighter to compensate for lugging heavier engines all the way to space, then you may actually lose a little payload performance.  If we were starting in orbit with 464.9s vs 452.3s of Isp, with full propellant tanks, then yes, you would see a meaningful payload performance increase, but again, it's not going to be a night-and-day type of improvement, especially if the nozzle extension mass is significant.

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But there is a problem here as far as estimating payload capability. Commonly with fixed nozzle stages you use a fixed value for the ISP to estimate the delta-v possible. This is convenient since it is only a single line calculation using the rocket equation. But that probably is not accurate for a variable nozzle since you are changing the characteristics of the nozzle through out the flight. You could you use, say, the vacuum isp for this case but that would probably over estimate it.

What really needs to be done is accurate trajectory simulations of the type NASA uses. Then you could see how the variable nozzle deviates from what a fixed nozzle would do. I’ve only seen one paper actually do that. I’ll see if I can find it.

  Bob Clark

Well,  they clearly have some sort of problem going on that they don't yet understand. 

The fact that something bad happened in a static test tied down to a thrust stand,  suggests a possible resonance in a feed plumbing line,  not a vehicle structural mode,  which would be strongly damped by the fixity of the thrust stand.  Such would explain a lot of the fires and engine outs we have been seeing all along.  But it takes a thrust or chamber pressure oscillation to do that,  one occurring at one of the organ pipe oscillation modes in that feed line. 

I really seriously doubt they have any sort of data acquisition that is not digital.  You cannot see an organized oscillation hidden in the noise "hash" with digitally-acquired data:  the inherent pixellation completely obscures it.  You need a high-speed FM tape recording using a recorder of about 1 MHz capability.  You play this back at various speeds and magnifications,  either as a paper printout or on a scope display,  until you find what you are looking for:  a definite periodicity with a nontrivial amplitude.  It will be a signal hidden among a nonperiodic "hash" of all sorts of amplitudes and frequencies,  so it may not initially "leap off the display" at you.  You do this with both thrust and pressure data recordings.

The reason I suspect they have a thrust/pressure oscillation going on is because of the 10-Hz shaking of my doors,  windows,  and hanging light fixtures when they test Raptors,  but NOT Merlins,  at the McGregor test facility.  You cannot hear this signal,  but you can feel and observe its effects.  10 Hz is below the lowest frequency humans can hear (but an elephant could).

This is the sort of problem where the basic approach of "build it,  break it,  build another" will let you down.  This is the sort of thing where you have to stop and investigate carefully just to figure out what went wrong.  SpaceX is not used to doing that.  SpaceX has no one on their payroll old enough to know about how to investigate anything like that.  And they have a bad "not invented here" culture that prevents them from learning from the mistakes of others,  especially mistakes made long ago.

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GW

This topic combines two ideas that are not normally associated.

Quantum pairing is thought to be superluminal so it would be interesting to see if anyone can find evidence supporting that conjecture.

Fusion is the focus of multiple topics already, but in ** this ** topic the idea of mechanical advantage is available if anyone would care to explore it.

Fusion is present in abundance in the Universe.  Every instance known to date is gravity mediated fusion.

(th)

That link in the quote in post 2060 just above does not take me to anything I ever sketched.  It takes me to the imgur webpage,  not to any particular image. 

Meanwhile,  just to let everybody know,  SpaceX has been testing a lot of Raptor engines lately,  quite a few up on the tower stand where they are quite loud.  There's about a 10-Hertz inaudible pressure wave from these tests that really rattles the windows and doors and chandelier-type light fixtures.  I feel it with Raptors,  but not Merlins.  I suspect the soot cloud in the flame with kerosene has an oscillation damping effect that is missing with the methane.

That too-low-to-hear but significant amplitude signal suggests a nontrivial thrust oscillation with a fairly-well-defined 10 Hertz frequency in the Raptor that the Merlins just do not have.  If there are any vehicle or plumbing modes near that 10-Hertz frequency,  resonance could easily cause POGO problems. 

Those tests have been running pretty close to 6-7 minutes each,  essentially a full duration ascent burn for either stage.  I cannot tell a vacuum Raptor from a sea level Raptor,  they all sound about the same,  from 6 miles away.  I suspect the vacuum tests are running with something approximating a normal shock,  essentially just barely aft of the exit plane.  They would have to be at near full pressure,  and would be separated between ignition at some reduced pressure,  and throttle-up to test conditions.   I usually don't hear the "explosions" that were full-power starts.  Von Braun found reduced-throttle ignition to be far more survivable,  about 80 years ago.
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GW