SSME based SSTO’s.

[sevenperforce]

13 hours ago, Exoscientist said:

Plus, you would then have a working SSTO. The payloads that the Falcon9 v1.1 had launched could be launched now by the first stage only, so for a lower cost not needing the upper stage.

Even F9v1.0 had an LEO payload of 9 tonnes. There’s no way that any modification of a F9B5 first stage would be able to loft anywhere near 9 tonnes as an SSTO. Maybe 1-2 tonnes max.

But if you only need to lift 1-2 tonnes to LEO to begin with, you’re not going to be buying a Falcon 9 flight; you’re gonna be getting a rideshare or going with a much cheaper smallsat launcher, so the comparison makes no sense.

And that’s without even factoring in reuse. It’s not lower-cost to throw away an entire reusable first stage just to avoid throwing away a much cheaper second stage. 

[Exoscientist]

 Rather than the commonly made statement that SSTO’s are not technically feasible or that they can’t carry significant payload, the actually truth of the matter is that if you use both high performance engines and lightweighted tanks, then SSTO’s can carry just as much or more payload in terms of payload fraction as do the current TSTO’s for expendable launchers.

 Moreover, because of the greatly reduced payload of the TSTO on boostback to launch site, the SSTO can actually carry more payload as a fully reusable launcher.

  That’s a stunning fact. The exact opposite of what is said about SSTO’s is the case.

     Robert Clark

Edited July 3, 2021 by Exoscientist
Clarity

[sevenperforce]

7 minutes ago, Exoscientist said:

if you use both high performance engines and lightweighted tanks, then SSTO’s can carry just as much or more payload in terms of payload fraction as TSTO’s for expendable launchers.

But this is simply not true.

It just isn’t. If your notional SSTO and notional TSTO both have access to the same high-performance engines and lightweight tanks, and both have the same GLOW, the TSTO will carry more payload to orbit. More as a fraction of GLOW and more as a fraction of total vehicle dry mass.

7 minutes ago, Exoscientist said:

because of the greatly reduced payload of the TSTO on boostback to launch site, the SSTO can actually carry more payload as a fully reusable launcher.

But it can’t. Even with boostback penalties, it can’t.

You can set engine performance, tank mass ratio, and heat shield/recovery mass **wherever** you want them, and a TSTO configuration will **always** carry more payload to orbit than an SSTO of equal GLOW. Without fail. 

[tater]

1 minute ago, Exoscientist said:

 Rather than the commonly made statement that SSTO’s are not technically feasible or that they can’t carry significant payload, the actually truth of the matter is that if you use both high performance engines and lightweighted tanks, then SSTO’s can carry just as much or more payload in terms of payload fraction as TSTO’s for expendable launchers.

 Moreover, because of the greatly reduced payload of the TSTO on boostback to launch site, the SSTO can actually carry more payload as a fully reusable launcher.

Most reusable TSTO flights don't use RTLS, they use ASDS landings downrange. And again, it does not matter. If a LV can loft more than your payload to the desired orbit, then the losses for reuse don't matter to the customer, it's not like SpaceX charges them MORE for reuse, they have lower launch costs.

 

1 minute ago, Exoscientist said:

  That’s a stunning fact. The exact opposite of what is said about SSTO’s is the case.

Payload fraction is irrelevant.

Cost. That's literally the only thing that matters.

An expended SSTO vs a TSTO that can be flown 10 times needs to be more than 10 times cheaper. As for stage 2 reuse that is cost effective, we have zero examples yet. Starship will be the first if it works. Saying it is easy, or will only require some small mass penalty is a statement without rigor. It might well be possible, but someone needs to actually do it, it's a complex problem that contains many known issues to solve, and some unknown ones.

[StrandedonEarth]

Well, I do see one potential benefit of an SSTO, especially with the RS-25's or RS-68's appetite for LH₂: wet labs. The large volume of hydrogen tanks is an advantage in that case, and not trying to re-enter the beast makes things much simpler. Sure, wet labs always seems to be too complicated, but I think the better approach is to design a hab that can function as a tank for awhile, versus trying to figure out how to convert a tank into hab space. But I digress, and that may be impractical for a tank the size of an LH2 SSTO tank. The engines can always be removed and dropped back dirtside.

Of course, if/when orbital manufacturing/recycling becomes a thing, SSTO tanks could quickly find a new life in space. But even then it could still be easier/better to convert than recycle.

[Exoscientist]

On 7/2/2021 at 8:32 AM, sevenperforce said:

Even F9v1.0 had an LEO payload of 9 tonnes. There’s no way that any modification of a F9B5 first stage would be able to loft anywhere near 9 tonnes as an SSTO. Maybe 1-2 tonnes max.

 Everyone knows the great importance of the Isp on the payload a rocket can carry to orbit. Imagine the Merlins used on the first stage having their vacuum Isp being raised from 312 s to 360+ s by using altitude compensation.

 Now also consider the weight saved on the aluminum-lithium tanks by using carbon fiber is about 30%. Not as good as the weight saved over standard grade aluminum of 50% but still pretty good.

 Try now the rocket equation with a vacuum Isp of 360+ s to see the payload the first stage can get orbit. You can use the dry mass and propellant mass numbers for the Falcon 9 v1.1 here: https://www.spacelaunchreport.com/falcon9ft.html#components

 You’ll be surprised by how high the payload is as an SSTO.

 Unfortunately, this isn’t quite accurate. The commonly used method of estimating the payload by the rocket equation is uncertain for the altitude compensation case. The reason is the estimate is known to be reasonably accurate for using fixed nozzle engines.  But its accuracy is unknown for the altitude compensation case.  A real trajectory simulation with an altitude compensating nozzle needs to be done to calculate this.

 

  Robert Clark

[Exoscientist]

On 7/3/2021 at 11:29 AM, sevenperforce said:

But this is simply not true.

It just isn’t. If your notional SSTO and notional TSTO both have access to the same high-performance engines and lightweight tanks, and both have the same GLOW, the TSTO will carry more payload to orbit. More as a fraction of GLOW and more as a fraction of total vehicle dry mass.

But it can’t. Even with boostback penalties, it can’t.

You can set engine performance, tank mass ratio, and heat shield/recovery mass **wherever** you want them, and a TSTO configuration will **always** carry more payload to orbit than an SSTO of equal GLOW. Without fail. 

  

You missed the point I’m making. The claim is that SSTO’s can’t be competitive to current rockets. But current rockets do not use max efficiency engines or max weight optimized stages!

 The importance of having a much higher Isp and weight optimized stages is so important that the SSTO will match or exceed the payload of the TSTO that do not use them.

 Yes, the payload of the TSTO that also does use them will be higher than the SSTO. But still the SSTO will be more efficient than the current rockets that do not.

 

  Robert Clark

[tater]

33 minutes ago, Exoscientist said:

You missed the point I’m making. The claim is that SSTO’s can’t be competitive to current rockets. But current rockets do not use max efficiency engines or max weight optimized stages!

 The importance of having a much higher Isp and weight optimized stages is so important that the SSTO will match or exceed the payload of the TSTO that do not use them.

 Yes, the payload of the TSTO that also does use them will be higher than the SSTO. But still the SSTO will be more efficient than the current rockets that do not.

 

It's a non-point. If there was any advantage in doing so, the reusable TSTO boosters would get the same treatment, lofting yet more payload.

Regardless, if the SSTO is not recoverable, and reusable with less effort than a stage 1, it's still a net loss. Throwing away a slightly better than current SSTO is still throwing an entire rocket away.

Your F9 example gets ~5t to LEO. Yeah, I just called it 360s for the whole flight. I did not reduce the dry mass at all, but if I did, it gets more like 13t to LEO. Now what? I have thrown away a F9 first stage to get pretty normal payloads to LEO. The alternative is to get the same payload to LEO—and recover the booster.

 

Edited July 6, 2021 by tater

[Rakaydos]

18 minutes ago, Exoscientist said:

 

 Yes, the payload of the TSTO that also does use them will be higher than the SSTO. But still the SSTO will be more efficient than the current rockets that do not.

 

  Robert Clark

so what are you trying to accomplish?

Because it looks like you want  to build a rocket with next gen tech, but aiming for numbers where real next gen  launchers will crush it, just for the claim to fame of being SSTO.

Sure. You can do that, if you have a few billion lying around that you dont need anymore. But you're not going to get any investers on board, because there's no return on that investment. 9 engine starship SSPTP is as close as you are going to get, and even that requires commonality with a heavy lift upper stage to make the buisness case close.

[Rhomphaia]

1 hour ago, Exoscientist said:

 Everyone knows the great importance of the Isp on the payload a rocket can carry to orbit. Imagine the Merlins used on the first stage having their vacuum Isp being raised from 312 s to 360+ s by using altitude compensation.

 

  Robert Clark

Except what is the point of increasing the mass and complexity to raise the vacuum efficiency of an engine that is never going to be used in a vacuum?

Staging is a form of altitude altitude compensation by itself, by limiting the range of altitude the rocket works under.

[tater]

Course the OP was about SSMEs. I want to know in what possible world the SSME can result in ANY cost-effective LV. All the dev money for the rest of the vehicle can be "free."

Do the economic math on how an SSME can work.

[Shpaget]

5 hours ago, Exoscientist said:

Imagine the Merlins used on the first stage having their vacuum Isp being raised from 312 s to 360+ s by using altitude compensation.

In case of F9, you can't fit 9 vacuum optimized Merlins at the bottom of F9. The nozzles are way to large.

 

[Exoscientist]

 

 For those interested in the topic of SSTO’s a nice book to read is:

Halfway to Anywhere: Achieving America's Destiny In Space.

https://www.amazon.com/Halfway-Anywhere-Achieving-Americas-Destiny/dp/0871318059/

 There, Stine also argues a SSTO can match the payload of a TSTO.

     Robert Clark

[Exoscientist]

2 hours ago, Shpaget said:

In case of F9, you can't fit 9 vacuum optimized Merlins at the bottom of F9. The nozzles are way to large.

 

 

 A fair point. One way would be to shorten the nozzles then place them around a central spike, as is planned for the Firefly rocket:

Robert Clark

[Rhomphaia]

5 hours ago, Exoscientist said:

 

 A fair point. One way would be to shorten the nozzles then place them around a central spike, as is planned for the Firefly rocket:

Robert Clark

As was planned for the firefly. ended in a typical fashion for aerospikes.

too heavy, too complex, too difficult to cool.

[RCgothic]

11 hours ago, Exoscientist said:

There, Stine also argues a SSTO can match the payload of a TSTO.

     Robert Clark

Then Stine is also wrong. Equating SSTO Vs TSTO for same DV:

ISPc*g*ln((P1+SSTOw)/(P1+SSTOd) = ISP1*g*ln((P2+2Sw+1Sw)/(P2+2Sw+1Sd)) + ISP2*g*ln((P2+2Sw)/(P2+2Sd))

Where:

ISP1 is 1st stage ISP.

ISP2 is 2nd stage ISP

ISPc is altitude compensating SSTO engine ISP with an average not exceeding ISP2.

P1 is SSTO payload.

P2 is TSTO payload.

SSTOw/d is SSTO wet/dry masses.

1Sw/d is TSTO first stage wet/dry masses.

2Sw/d is TSTO second stage wet/dry masses.

This has no solutions where the SSTO payload P1 is greater than the TSTO payload P2 for similar gross lift off weight and similar optimisation of engine performance and tank fraction.

[DDE]

1 hour ago, RCgothic said:

Equating SSTO Vs TSTO for same DV:

I suspect the argument is that a much bigger SSTO can match the payload of a TSTO...

[Exoscientist]

2 hours ago, RCgothic said:

Then Stine is also wrong. Equating SSTO Vs TSTO for same DV:

ISPc*g*ln((P1+SSTOw)/(P1+SSTOd) = ISP1*g*ln((P2+2Sw+1Sw)/(P2+2Sw+1Sd)) + ISP2*g*ln((P2+2Sw)/(P2+2Sd))

Where:

ISP1 is 1st stage ISP.

ISP2 is 2nd stage ISP

ISPc is altitude compensating SSTO engine ISP with an average not exceeding ISP2.

P1 is SSTO payload.

P2 is TSTO payload.

SSTOw/d is SSTO wet/dry masses.

1Sw/d is TSTO first stage wet/dry masses.

2Sw/d is TSTO second stage wet/dry masses.

This has no solutions where the SSTO payload P1 is greater than the TSTO payload P2 for similar gross lift off weight and similar optimisation of engine performance and tank fraction.

 Again, when I say an “SSTO can match or beat a TSTO in payload”, that is shorthand for this:

Current rockets do not use maximally efficient engines or weight optimized structures on their first stages. But high Isp and low weight are extremely important for maximizing a rockets payload to orbit. So  when you do use both of these, it radically improves payload to LEO. The result is the SSTO can match or exceed the payload of a current TSTO. And the two stage version of such an optimized rocket can double the payload of a current rocket.

  Robert Clark

1 hour ago, DDE said:

I suspect the argument is that a much bigger SSTO can match the payload of a TSTO...

 No, I’m not. I am literally saying an SSTO that uses both maximum Isp engines and weight optimized structures can exceed in payload fraction that of a current TSTO, which do not use either of these.

 So the same size optimized SSTO can exceed the payload of an unoptimized TSTO.

 

    Robert Clark

[DDE]

24 minutes ago, Exoscientist said:

So the same size optimized SSTO can exceed the payload of an unoptimized TSTO.

And so you admit defeat, since your best comparison is apples to oranges.

[Beccab]

1 hour ago, Exoscientist said:

So the same size optimized SSTO can exceed the payload of an unoptimized TSTO.

...which means that SSTO don't have actually a reason to exist? Spending a huge amount of money to create a perfectly optimised SSTO would give you actually less payload to orbit, for more complexity and higher cost for refurbishment (since in a TSTO most of the rocket reenters before reaching orbital velocity, requiring no heat shield while in an SSTO the whole rocket has to resist reentry and be exposed to the plasma). But hey, if you stop advancing the technology used in every other kind of rocket, eventually it could be comparable!

[Rakaydos]

15 hours ago, Exoscientist said:

 

 A fair point. One way would be to shorten the nozzles then place them around a central spike, as is planned for the Firefly rocket:

Robert Clark

I suggest you watch this video: https://everydayastronaut.com/aerospikes/

Rocketlab already uses gimbal to get some aerospike-like effects out of electron. But actual aerospikes are just not worth the effort.

[sevenperforce]

23 hours ago, Exoscientist said:

Imagine the Merlins used on the first stage having their vacuum Isp being raised from 312 s to 360+ s by using altitude compensation.

How in the world would a sea-level Merlin 1D get 360 seconds or more of specific impulse via “altitude compensation” when the actual Merlin 1D Vacuum only gets 348 seconds??

Plugging in nonsense numbers yields nonsense answers. 

2 hours ago, Exoscientist said:

an SSTO that uses both maximum Isp engines and weight optimized structures can exceed in payload fraction that of a current TSTO, which do not use either of these.

But (if true) how is this relevant? If you have access to “maximum Isp engines” and “weight optimized structures” then just use them to build a smaller TSTO that delivers your desired payload more cheaply than an SSTO.

[RCgothic]

Well there probably is additional energy that *could* be extracted from kerolox using some monster full flow stage combustion cycle or similar.

I don't have the numbers to hand, but it'd be a similar argument that the theoretical maximum for Hydrolox is around 530s for a stochiometric mix, so the 450s of RS25 and 465s of RL10 aren't 100% efficient.

There are good reasons the theoretical maximum isn't *practically* acheivable, such as the 9GW required to turn the turbopumps in RS25, and back-pressure from the atmosphere.

But suffice it to say, any advanced SSTO with technology that could put it on a par with *current* TSTO would be at least one and probably both of the following:

1) More expensive than a current TSTO due to advanced manufacturing techniques.

2) Incapable of matching a similarly advanced TSTO design.

Edited July 7, 2021 by RCgothic

[sevenperforce]

On 6/24/2021 at 4:42 AM, Exoscientist said:

if I were to actually make an SSTO out of the Delta IV I would probably just add altitude compensation to the RS-68 already on the first stage. This way you could even get the vacuum Isp to the 465+ s range.

Missed this before, but the RS-68 is a gas generator cycle. You're not going to get anywhere near 465 seconds or higher...more like the YF-75's 438 seconds, if that. The RS-68 already gets 430 seconds of vacuum specific impulse.

[sevenperforce]

6 hours ago, RCgothic said:

Well there probably is additional energy that *could* be extracted from kerolox using some monster full flow stage combustion cycle or similar.

The RD-0124 uses oxygen-rich staged combustion to get a whopping 359 seconds of vacuum specific impulse, but it does so at the expense of an engine which (a) cannot be fired at sea level, (b) cannot be throttled, and (c) has a rather poor T/W ratio of only 52.5 at its highest-thrust configuration.

You cannot go full-flow staged combustion with kerolox because of coking issues, so 359 seconds is probably the best you're going to get, and of course you don't get that just by slapping a different nozzle on the Merlin 1D.

The RS-25 and YF-90 are also staged-combustion engines which share a propellant type but the RS-25's altitude-compensating nozzle means it loses about 2 seconds of specific impulse compared to the vacuum-optimized YF-90. It also has a sea level specific impulse 19% lower, which also means sea level thrust that is 19% lower.

So if you tried to build an SSTO using a version of the RD-0124 with an altitude-compensating nozzle, you'd end up with a sea level specific impulse of around 289 seconds, a vacuum specific impulse of 357 seconds, and a sea level T/W ratio of only 42.1, which means you're REALLY gonna have trouble getting off the ground. SSTOs need high thrust at launch...you wouldn't want to go any lower than 1.2 at the very minimum. So your engine is going to be at least 2.85% of your GLOW.

If you're aiming to match or exceed a 3% payload fraction (as @Exoscientist said upthread), then that means nearly 6% of your mass is just payload and engine, nothing else. Let's also assume, just to simplify the math, that the first 30% of the burn takes place at the fixed sea-level specific impulse of 289 seconds and the latter 70% of the burn takes place at the fixed vacuum specific impulse of 357 seconds. Next, let's assume, ad arguendo, that your tanks literally weigh **nothing**.

m₀ = 1, m₁ = 0.718, m_f = 0.0585

dV₁ = 939 m/s, dV₂ = 8778 m/s, dV_f = 9.72 km/s

That's barely enough to get into a usable orbit, and probably won't be once you factor in the higher gravity drag from your lower T/W ratio. And we haven't accounted for the weight of actual propellant tanks. The tanks and structure of the Atlas D (not including engines) equaled 4.2% of the total propellant mass. Even if "carbon composites" can magically cut that by 40% to just 2.37% of GLOW, you've already gone from a 3% payload fraction to an 0.63% payload fraction.

And for what? Just so you can say you built an SSTO? What possible reason could there be to not simply stage it and deliver 5X as much payload for the same GLOW?