SSME based SSTO’s.

[Exoscientist]

17 hours ago, DDE said:

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

 Actually, not. This shows why both these advances should be undertaken. By using both, we can reduce the cost of space and improve on the payloads of rockets we have now.

   Robert Clark

[Exoscientist]

10 hours ago, sevenperforce said:

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.

 Yes. But for an upper stage getting high vacuum Isp is largely an effect of using a large expansion ratio. For instance the Merlín Vacuum at 348 s vacuum Isp has a expansion ratio of 164 to 1 using a quite large nozzle attachment:

 

 But the Russian RD-58s gets a vacuum Isp of 361 s by using an expansion ratio of 189 to 1:

https://web.archive.org/web/20160407043338/http://www.friends-partners.org/partners/mwade/engines/rd58s.htm

 The importance of altitude compensation is it allows you to get the very high vacuum Isp of a upper stage engine while being being able to launch from the ground, as a sea level engine. It does this by using a variable area nozzle, or a nozzle such as the aerospike that acts as a variable area nozzle.  

 Actually the 348 s Isp of the Merlin Vacuum is quite good. Even if you used a variable area nozzle on the sea level Merlins that just reached the 348 s value, that would also greatly increase the TSTO payload, and allow an SSTO with significant payload.

 

  Robert Clark

Edited July 8, 2021 by Exoscientist

[DDE]

1 hour ago, Exoscientist said:

 Actually, not. This shows why both these advances should be undertaken. By using both, we can reduce the cost of space and improve on the payloads of rockets we have now.

   Robert Clark

Poitnless redundancy leads to increased costs.

[Exoscientist]

10 hours ago, sevenperforce said:

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.

 

 The vacuum Isp of the RS-68 is actually only 412 s. That is why performance would be radically improved by increasing this to 470 s. You can get a high vacuum Isp on an upper stage engine just by using a nozzle extension. So for example the RL-10 engine on the Centaur upper stage gets a  ca. 462 to 465.5 s Isp by using an extendable nozzle attachment.

 

  Robert Clark

[Rakaydos]

2 hours ago, Exoscientist said:

 Actually, not. This shows why both these advances should be undertaken. By using both, we can reduce the cost of space and improve on the payloads of rockets we have now.

   Robert Clark

You haven't shown how ssto will reduce the cost of space, and we've all demonstrated that the same effort will improve the payloads of two-stage to orbit rocket better than ssto. Chemical ssto is an always will be pointless on earth.

[RCgothic]

4 minutes ago, Exoscientist said:

 

 The vacuum Isp of the RS-68 is actually only 412 s. That is why performance would be radically improved by increasing this to 470 s. You can get a high vacuum Isp on an upper stage engine just by using a nozzle extension. So for example the RL-10 engine on the Centaur upper stage gets a  ca. 462 to 465.5 s Isp by using an extendable nozzle attachment.

 

  Robert Clark

The extendable nozzle on some variants of RL10 isn't for altitude compensation, it's to reduce interstage length.

 

Merlin 1D has an expansion ratio of just 16. It is the best kerolox booster engine ever designed. It has 7-19s more ISP than the F1, twice the TWR and nearly twice the thrust per unit area.

If you slap a larger nozzle on it, not only would flow separation destroy it, but it would weigh more, have lower thrust per unit area, and,  crucially, no longer fit in a cluster under an F9.

 

Similarly RS25 has an expansion ratio of 69. To have an expansion ratio of 189 to extract full ISP it would need a significantly larger, heavier nozzle. As a hydrolox engine it already has woeful thrust for a booster engine (hence the SRBs), and the larger nozzle (assuming it survives flow separation issues) would cut its thrust per unit area by approx 2.5x and also reducing the number of engines you can fit under a notional stage by 2.5 times.  RS25 SSTO isn't getting off the pad.

[Hannu2]

On 7/3/2021 at 6:19 PM, 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 do the current TSTO’s for expendable launchers.

 

Current rockets use very optimal tanks and engines. Every company develops their products continuously. If you take some values from another products, test articles or even theoretical values, they are scifi speculation and not facts. If someone knew how to use such materials and components in orbital boosters they certainly would use them (in two stage rockets to maximize payload per cost).

 

On 7/3/2021 at 6:19 PM, Exoscientist said:

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

 

Comparing existing real tech to some speculative values is very useless fact even it is formally true statement.

[Hannu2]

On 7/2/2021 at 1:51 AM, Exoscientist said:

 More precisely, it did a trade that for the Starship when you take into account the weight of the thermal shielding  the ultra strong steel it decided on would be more heat resistant so less thermal shielding would be required.  Carbon fiber tanks are now a well-established technology.

Carbon fiber is not established in large rocket booster tanks. It is very special thing which needs special properties. If there are high pressure carbon fiber gas cylinders they are not straightly usable in rockets. That SpaceX's "trade" means failure. Their composite tanks were not good enough to heavy booster and they did not see way to achieve objectives in predictable time and costs. There are also no other companies who use such tanks in their large rockets. SpaceX use steel and Li-Al and others Li-Al.

That 304L is not "ultra strong space steel with 10 % unobtainium and cost of million per kg" but very basic stainless steel used in very many industrial applications. Especially in food industry because it is very safe. Also usual kitchen stuff, like forks, spoons and pots, are made from such steel. They said their will develop material later phases and I do not know are new SNs still made from commercial 304L steel.

 

[sevenperforce]

9 hours ago, Exoscientist said:

 Yes. But for an upper stage getting high vacuum Isp is largely an effect of using a large expansion ratio.

No, peak Isp is a combination of three main things: propellant choice/mixture ratio, power cycle/chamber pressure, and expansion ratio. All three are significant.

Go take a look at the comparison of orbital rocket engines page on Wikipedia. I just added a column for all power cycles so you can make comparisons. Power cycles are EXTREMELY important.

9 hours ago, Exoscientist said:

For instance the Merlín Vacuum at 348 s vacuum Isp has a expansion ratio of 164 to 1 using a quite large nozzle attachment:

<snip>

 But the Russian RD-58s gets a vacuum Isp of 361 s by using an expansion ratio of 189 to 1:

<snip>

The RD-58 (and virtually all Russian or Russian-derived kerolox engines) achieve high vacuum specific impulse by using staged combustion. The staged combustion cycle is more efficient than the gas generator cycle. You will NEVER be able to reach staged-combustion levels of specific impulse merely by adding a larger nozzle to a gas generator engine.

9 hours ago, Exoscientist said:

Actually the 348 s Isp of the Merlin Vacuum is quite good. Even if you used a variable area nozzle on the sea level Merlins that just reached the 348 s value, that would also greatly increase the TSTO payload, and allow an SSTO with significant payload.

You would not be able to reach 348 seconds of vacuum specific impulse by adding an altitude-compensating nozzle to a sea-level Merlin 1D engine. At most you could reach around 346 seconds. Additionally, your engine would weigh almost twice as much, and it would be much larger.

If SpaceX could simply snap their fingers and add an altitude-compensating nozzle to the Falcon 9 lower stage to increase TSTO, why wouldn't they have done so already? It doesn't make sense.

In reality, the Merlin 1D already uses an altitude-compensating nozzle. It is overexpanded at sea level, just like the RS-25. If SpaceX made it any bigger, they wouldn't be able to fit as many of them underneath the Falcon 9, so they would lose payload.

9 hours ago, Exoscientist said:

The vacuum Isp of the RS-68 is actually only 412 s.

My mistake, I was thinking of the YF-77. You are correct that the vacuum specific impulse of the RS-68A is 412 seconds.

9 hours ago, Exoscientist said:

That is why performance would be radically improved by increasing this to 470 s.

Yes, the performance of the RS-68 would be radically improved if it had a vacuum specific impulse of 470 seconds. You could achieve this by replacing the RS-68 with a completely different engine that is capable of reaching 470 seconds. The RS-68 is not capable of reaching 470 seconds of vacuum specific impulse, no matter how ridiculously large of a nozzle extension you give it. It is a gas generator engine and so it will never be able to exceed 440-450 seconds. Consider, for reference, the LB-5 engine, which is also hydrolox and is also a gas generator. It already has a vacuum nozzle and it pulls just under 450 seconds.

9 hours ago, Exoscientist said:

You can get a high vacuum Isp on an upper stage engine just by using a nozzle extension.

You can get higher vacuum specific impulse on an engine by giving it a larger nozzle extension. However, this produces diminishing returns. You cannot get more energy out of propellants than the engine power cycle provides them. There is no free lunch.

9 hours ago, Exoscientist said:

So for example the RL-10 engine on the Centaur upper stage gets a  ca. 462 to 465.5 s Isp by using an extendable nozzle attachment.

The RL-10 is a closed expander cycle. Of course it gets a higher peak specific impulse than an engine that dumps turbine exhaust through a gas generator.

By analogy: wearing brass knuckles makes punching do more damage. But wearing brass knuckles does not make you punch harder. No matter how large a set of brass knuckles you give to a ten-year-old, they will not be able to punch as hard as Mike Tyson. They will do more damage than if they didn't have the brass knuckles, but they still have a limit.

[Exoscientist]

On 7/8/2021 at 5:11 AM, RCgothic said:

The extendable nozzle on some variants of RL10 isn't for altitude compensation, it's to reduce interstage length.

Merlin 1D has an expansion ratio of just 16. It is the best kerolox booster engine ever designed. It has 7-19s more ISP than the F1, twice the TWR and nearly twice the thrust per unit area.

If you slap a larger nozzle on it, not only would flow separation destroy it, but it would weigh more, have lower thrust per unit area, and,  crucially, no longer fit in a cluster under an F9.

Similarly RS25 has an expansion ratio of 69. To have an expansion ratio of 189 to extract full ISP it would need a significantly larger, heavier nozzle. As a hydrolox engine it already has woeful thrust for a booster engine (hence the SRBs), and the larger nozzle (assuming it survives flow separation issues) would cut its thrust per unit area by approx 2.5x and also reducing the number of engines you can fit under a notional stage by 2.5 times.  RS25 SSTO isn't getting off the pad.

 

 The designers of the Centaur want the long nozzle because they do want the high Isp in vacuum it provides. They could get this just in being fixed. But by making it extendible they are able to fit it in a shorter interstage. 

This is a key principle about rocket engines that by using wider nozzles you are able to get higher vacuum Isp. The problem with using  it at sea level is such large nozzles cause what is called flow separation. This is a dangerous instability condition that can literally tear an engine apart. This is why such large nozzles are not used at sea level.

 The idea behind variable area nozzles is they are small at sea level and extend larger at altitude in vacuum. This effect can be emulated also by using an aerospike. If we are to make it be an actual extensible nozzle we can use recent materials using ceramics that can cut the weight by a factor of 3.

 

    Robert Clark

On 7/8/2021 at 9:10 AM, Hannu2 said:

Carbon fiber is not established in large rocket booster tanks. It is very special thing which needs special properties. If there are high pressure carbon fiber gas cylinders they are not straightly usable in rockets. That SpaceX's "trade" means failure. Their composite tanks were not good enough to heavy booster and they did not see way to achieve objectives in predictable time and costs. There are also no other companies who use such tanks in their large rockets. SpaceX use steel and Li-Al and others Li-Al.That 304L is not "ultra strong space steel with 10 % unobtainium and cost of million per kg" but very basic stainless steel used in very many industrial applications. Especially in food industry because it is very safe. Also usual kitchen stuff, like forks, spoons and pots, are made from such steel. They said their will develop material later phases and I do not know are new SNs still made from commercial 304L steel.

 

 Actually the grade SpaceX wants to use is much stronger than standard steel and much more expensive. As I mentioned in my blog post there are some metals now that can match and exceed carbon composites. That’s perfectly fine, as long as they reduced the structural mass of the stage.

 Carbon composites stages are used in some rocket stages, most commonly in solid rockets. SpaceX didn’t use it not because it wouldn’t work, but because the carbon composites are far less heat resistant than steel and would have required more mass for thermal shielding.

 

  Robert Clark

[Exoscientist]

 We’ve been discussing the advantages or disadvantages of making such advances such as max Isp engines and max lightened stages. This has only been estimated by the rocket equation. What really needs to be done is a Kerbal  simulation using the Real Solar System mod or Realism Overhaul to show the Delta IV given SSME’s and lightweighted stages really can get the high payloads as an SSTO and TSTO suggested by the rocket equation estimates.

  Robert Clark

[Exoscientist]

On 7/8/2021 at 2:09 PM, sevenperforce said:

The RD-58 (and virtually all Russian or Russian-derived kerolox engines) achieve high vacuum specific impulse by using staged combustion. The staged combustion cycle is more efficient than the gas generator cycle. You will NEVER be able to reach staged-combustion levels of specific impulse merely by adding a larger nozzle to a gas generator engine.

 I’m sorry but this is simply incorrect. Any rocket engine can get high vacuum Isp by using a long extension nozzle, as indicated by the Merlin Vacuum getting 348s Isp.  The expander cycle used on the RL-10 is actually one of the least efficient cycles used on rocket engines. It was used on the RL-10 because as an upper stage engine all it needed was a long nozzle, not combustion efficiency, to get high vacuum Isp.

 Remember this is in regards to vacuum Isp. For sea level engines you need high thrust and gas generator cycles and staged combustion cycles are better at providing that.

 

    Robert Clark 

Edited July 9, 2021 by Exoscientist

[DDE]

12 hours ago, Exoscientist said:

The expander cycle used on the RL-10 is actually one of the least efficient cycles used on rocket engines. It was used on the RL-10 because as an upper stage engine all it needed was a long nozzle, not combustion efficiency, to get high vacuum Isp.

You're simply wrong.

Quote

The expander-engine cycle relies on using a cryogenic fuel, which is gasified and heated in the TC cooling jacket, to drive the turbine(s). The relatively cool turbine exhaust gas of evaporated hydrogen is subsequently fed into the combustion chamber. There are no GGs or preburners. Figure 4.2-7 shows the simple flow diagram of the first such LPRE with an expander cycle. The performance of such an engine is slightly better (2-7% depending on design details) than the gas-generator cycle, but the internal fuel pressures and inert engine mass are somewhat higher than an engine with an equivalent GG cycle. There is no real performance penalty when compared to the staged combustion cycle, which is discussed next. This expander cycle works only with a cryogenic fuel that can be evaporated, such as hydrogen. It would not work with storable fuels, such as kerosene or UDMH. To date all LPREs with an expander engine cycle have used LOX/LH2. The RL-10 engine of Pratt & Whitney was the first in the world to use the expander cycle for an upper-stage launch-vehicle application. Its flow diagram is in Fig. 4.2-7. It was first tested on the ground in 1959 and flew first in 1963. Improved and uprated versions of the RL-10 continue to fly using this cycle. The Russians built and tested an experimental engine with an expander cycle in 1998, and Rocketdyne developed one between 1981 to 1983, but none of these LPREs have flown. 

...

The staged combustion cycle allows a higher performance LPRE (2 to 7% better than a gas-generator cycle LPRE, depending on the design) and about the same performance as a LPRE with an expander-engine cycle.

- George P. Sutton, History of Liquid Propellant Rocket Engines

The only reason everyone doesn't go staged combustion-only - as the Soviets liked to do, they even had staged-combustion verniers on SLBMs - is that it's often too much fuss for too little gain.

Much like SSTOs.

Edited July 10, 2021 by DDE

[Exoscientist]

4 hours ago, DDE said:

but the internal fuel pressures and inert engine mass are somewhat higher than an engine with an equivalent GG cycle.

 This part in the passage you quoted from Sutton that you did not highlight is why expander cycle engines are not used for sea level engines and gas generator and staged combustion engines are used instead, because these types of engines can provide more power, which is needed for launch from the ground:

Liquid Rocket Propulsion.

Engine Cycles – Expander

•  Relies on a turbopump to force propellants from tanks to the combustor

•  Tanks kept at lower pressures

•  Fuel heated via regenerative cooling process and passed through turbine to drive pumps

•  Thrust-limited due to square- cube rule (heat transfer)

•  RL10 (Delta IV, Atlas V), LE-5B (H-IIA, H-IIB)

http://rocket.gtorg.gatech.edu/files/slides/Liquid_Rocket_Propulsion.pdf

 In any case, it is still the case that ANY engine of any combustion cycle can get high vacuum Isp by using long nozzles.

  Some further references:

Rocket Engine Nozzle

https://en.wikipedia.org/wiki/Rocket_engine_nozzle

 

Rocket Propulsion Elements, by Sutton and Biblarz

http://mae-nas.eng.usu.edu/MAE_5540_Web/propulsion_systems/subpages/Rocket_Propulsion_Elements.pdf  [full text]

 

Rocket Propulsion.

http://www.braeunig.us/space/propuls.htm

 

 Robert Clark

Edited July 10, 2021 by Exoscientist

[DDE]

14 minutes ago, Exoscientist said:

This part in the passage you quoted from Sutton that you did not highlight is why expander cycle engines are not used for sea level engines and gas generator and staged combustion engines are used instead, because these types of engines can provide more power, which is needed for launch from the ground

That wasn't your argument:

17 hours ago, Exoscientist said:

The expander cycle used on the RL-10 is actually one of the least efficient cycles used on rocket engines.

And the broader issue under discussion was whether a greater nozzle expansion ratio can save an otherwise inferior engine. It cannot.

[Exoscientist]

15 minutes ago, DDE said:

That wasn't your argument:

And the broader issue under discussion was whether a greater nozzle expansion ratio can save an otherwise inferior engine. It cannot.

 

 The expander cycle is among the least efficient engines for sea level engines, due to its limited thrust. That is why it is used for upper stage engines where high thrust is not required. But a gas generator and staged combustion engine used for sea level launch can also get high vacuum Isp by using an extended nozzle. 

 

  Robert Clark

[sevenperforce]

16 hours ago, Exoscientist said:

The expander cycle used on the RL-10 is actually one of the least efficient cycles used on rocket engines.

Ah, yes. The expander cycle. Famously inefficient. So inefficient, in fact, that the RL-10A-3 used on the Saturn I S-IV Second stage could only develop 444 seconds of specific impulse.

Without a nozzle extension.

16 hours ago, Exoscientist said:

Any rocket engine can get high vacuum Isp by using a long extension nozzle, as indicated by the Merlin Vacuum getting 348s Isp.

Yes, you can achieve higher vacuum efficiency by using a longer nozzle, but that does not mean that any engine can achieve arbitrarily high specific impulse by using an arbitrarily long nozzle.

The M1DVac is already nearly maxed out. If you doubled or even tripled the length of the nozzle, it still would not be anywhere near the 360 seconds that the RD-0124 gets. You could get 350, maybe 351 if you were really lucky, but that is it. The energy simply isn’t there.

Similarly, adding an arbitrarily long nozzle extension to an RS-68 would not achieve arbitrarily high specific impulse. You might be able to get up to 437-439 seconds but you  would never be able to achieve the 444 seconds that the RL-10 gets without any nozzle extension at all.  You certainly couldn’t get the 460-470 seconds that the RL-10 achieves with a proper nozzle extension.

Quote

 Remember this is in regards to vacuum Isp. For sea level engines you need high thrust and gas generator cycles and staged combustion cycles are better at providing that.

You do need high thrust for sea level engines, but you are confused with respect to the way that gas generator engines and stage combustion engines achieve that.

The specific impulse of any rocket engine depends on just two things: the amount of energy imparted to the exhaust and the average molecular weight of that exhaust. However, the amount of energy imparted to the exhaust, in turn, depends on the following sequence:

• The specific energy of the propellants
• The percentage of the propellant’s specific energy which is burned in the combustion chamber
• The percentage of the propellant which passes through the combustion chamber
• Total combustion efficiency
• The pressure in the combustion chamber
• The pressure at the nozzle exit

All of these factors are important. You cannot ignore them. Adding a longer nozzle extension merely lowers the pressure at the nozzle exit, the very last step in the process. And while that is definitely very important, it cannot compensate for losses earlier in the process.

Up until the last two variables, chamber pressure and exit pressure, a closed expander cycle engine is always the most efficient thermodynamic cycle because 100% of the specific energy is extracted in the combustion chamber and 100% of the propellant passes through the combustion chamber. A staged combustion engine is slightly less efficient because part of the specific energy is lost to the preburner(s), even though 100% of the propellant still passes through the chamber. A gas generator cycle is the least efficient of the three because part of the specific energy is lost in the gas generator and part of the propellant is lost in the gas generator exhaust.

The hidden variable is the chamber pressure. The bigger the pressure drop, the more of the specific energy that gets converted into thrust. For a sea level engine, exit pressure at the nozzle is limited by atmospheric pressure, so you need your chamber pressure to be as high as possible so that most of the energy is converted into thrust before you reach 1 bar. Gas generator engines and staged combustion engines are good at this because they have a more energetic thermodynamic cycle, which gives them more available work potential to push the propellants into the chamber at the maximum possible pressure.

That is why simply adding a larger nozzle to a sea-level gas generator engine cannot make it competitive with a full-sized nozzle on a closed expander cycle engine. The gas generator has already sacrificed efficiency by burning some of the propellant in the gas generator and dumping the exhaust overboard, and it has already extracted most of the remaining energy from its propellants before its main exhaust reaches one bar. There’s just not that much energy left.

[sevenperforce]

8 minutes ago, Exoscientist said:

The expander cycle is among the least efficient engines for sea level engines, due to its limited thrust. That is why it is used for upper stage engines where high thrust is not required. But a gas generator and staged combustion engine used for sea level launch can also get high vacuum Isp by using an extended nozzle. 

A closed expander cycle engine is somewhat less efficient for sea level purposes not so much because of low thrust, but because of low chamber pressure. Low chamber pressure means less of a pressure drop between chamber pressure and atmospheric pressure, which leads to wasted potential.

Likewise, it is preferred for upper stages not so much because those stages don’t need as much thrust, but because operating in a vacuum means a larger nozzle can extract that wasted potential.

Adding a larger vacuum nozzle to a gas generator engine will extract whatever wasted potential remains, but it cannot match the vacuum specific impulse of a closed expander cycle engine because it does not have as much remaining potential energy as a closed expander cycle engine.

No matter how large of a nozzle you put on an RS-68, it will not be able to get anywhere close to what an RL-10 can produce with even a relatively small nozzle extension.

[tater]

1 hour ago, sevenperforce said:

No matter how large of a nozzle you put on an RS-68, it will not be able to get anywhere close to what an RL-10 can produce with even a relatively small nozzle extension.

And RS-68s cost ~$20M each.

This entire thread is predicated on SSME SSTOs being a thing, and cost effective. Using any SSMEs at all instantly makes pretty much any design "not cost effective" as a starting point as the cost—completely ignoring billions spent on just the recent SLS-era dev—is $100M each (supposedly simplified for expendable use, so this is low for a reusable one). This means to fulfill the OP claim, he needs to demonstrate a plausible, 100% reusable SSME based SSTO, and it needs to operate with effectively no refurb costs, since the initial ante on engines alone will be large. A single SSME based SSTO could get the pro-rata engine cost down to $10M with 10 launches with zero refurb cost. Given that the Shuttle-era SSMEs used for SLS cost $127M each to refurb, zero cost refurb seems... optimistic.

In short the OP claim is a non-starter.

Moving on to alternate engines is a change from the initial claim, and the dev from AJR also makes it a non-starter. AJR sells crazy expensive engines, not reasonably priced engines, and certainly not super low cost engines. That is unlikely to change as they are now going to be part of LockMart.

Be-4 (assuming they ever get it ready) would be a decent choice for a reusable, but the I_sp will be limited to what is possible for CH₄ with that engine cycle (nice addition to the wiki, @sevenperforce). We know the cost is ~$7M each (retail to ULA per Tory Bruno rocket cost math). So maybe that's a better starting point for messing with SSTO designs that are reusable. Of course you cannot change the engine, or you need to dev a new one, so it's use the off the shelf Be-4, or it's fantasyland again.

 

[StrandedonEarth]

It all makes me wonder: If someone wanted to redesign the RS-25 for the lowest manufacturing cost, how would it turn out? Assume that means that it stays hydrolox, despite that probably not being the best choice for SSTO, unless wetlabbing  it…

How low could the incremental price go, ignoring dev and factory cost? Would there be much performance hit? If only the entire thing could be done with additive manufacturing. Not a silver bullet; slow, I know, but still faster than some methods. 

Maybe returning a hydrolox SSTO wouldn’t be so hard; it may be so ‘fluffy’ that it slows down fast enough to not melt without a lot of protection. Did NASA ever do any experiments with ET reentries, seeking how long they could delay the breakup? Probably not because they wanted as much destruction of the tank as possible. 

Edited July 10, 2021 by StrandedonEarth

[tater]

52 minutes ago, StrandedonEarth said:

It all makes me wonder: If someone wanted to redesign the RS-25 for the lowest manufacturing cost, how would it turn out? Assume that means that it stays hydrolox, despite that probably not being the best choice for SSTO, unless wetlabbing  it…

How low could the incremental price go, ignoring dev and factory cost? Would there be much performance hit? If only the entire thing could be done with additive manufacturing. Not a silver bullet; slow, I know, but still faster than some methods. 

Maybe returning a hydrolox SSTO wouldn’t be so hard; it may be so ‘fluffy’ that it slows down fast enough to not melt without a lot of protection. Did NASA ever do any experiments with ET reentries, seeking how long they could delay the breakup? Probably not because they wanted as much destruction of the tank as possible. 

Build a rocket engine factory that turns out 2-4 in a day vs that many in a year?

The SSME is a great engine, but I don't think it should cost more than Be-4 (~$7M retail), or Raptor (unknown cost, maybe internal SpaceX cost of ~$1-2M with an aspirational internal cost target of $250k). Both engines are higher thrust, BTW, albeit with methane as the propellant, so lower I_sp. Unsure if propellant conversion would be possible, but the LR87 burned N2O4/Aerozine-50, and was converted to multiple other propellants, including hydrolox.

 

 

[sevenperforce]

2 hours ago, StrandedonEarth said:

It all makes me wonder: If someone wanted to redesign the RS-25 for the lowest manufacturing cost, how would it turn out? Assume that means that it stays hydrolox, despite that probably not being the best choice for SSTO, unless wetlabbing  it…

How low could the incremental price go, ignoring dev and factory cost? Would there be much performance hit? If only the entire thing could be done with additive manufacturing. Not a silver bullet; slow, I know, but still faster than some methods. 

There are a number of things you could do to simplify the design somewhat. Small performance hit but you could probably get better thrust out of it.

The RS-25 uses two fuel-rich preburners: one for the LOX pump turbine shaft and one for the fuel pump turbine shaft. Instead, you could just have a single preburner and split the exhaust out to two separate turbines, like the gas generator exhaust on the RS-68 and YF-77. Slight efficiency loss because the preburner exhaust would have a longer path to the turbines, but only using a single preburner would probably save weight anyway.

Or, hear me out.......

Expander cycle engines are simple and reliable, but they have poor thrust-to-weight and relatively low chamber pressure because there is only so much heat you can extract from the engine bell, and that gets worse and worse as your engine gets bigger.

However, what if you did a combination? Keep the same fuel-rich preburner for the LOX turbopump, but use a closed expander cycle for the fuel turbopump. Use the additional heat from the preburner to increase the thermodynamic potential in the expander cycle loop.

The LH2 would flow around the nozzle, chamber, AND preburner in order to absorb the maximum amount of heat. It would run through a turbine (which would turn the LH2 turbopump shaft), then exhaust into the preburner, which would burn with a small amount of LOX. The fuel-rich exhaust would run through another turbine (which would turn the LOX turbopump shaft) and then exhaust into the chamber.

The highest pressure would be in the LH2 coolant loop, which is an extremely gentle thermodynamic cycle in comparison to a preburner cycle. You still need a seal between the preburner turbine shaft and the LOX pump but that’s not nearly as challenging as dealing with an oxidizer-rich preburner. And since the preburner is only running the LOX turbopump, it is smaller and uses less energy than if you had one preburner for both pumps or two separate preburners.

Another approach, if you’re looking for a proper SSTO engine (for whatever reason), would be a variable-mixture tripropellant engine running LH2 *and* methane with LOX. The thermodynamic engine cycle for **that** would be insanity, though. Maybe you could do a hydrolox gas generator with exhaust to a LOX turbine and a fuel turbine, geared to both a LH2 pump and a methane pump. LH2 and methane play well enough together that the whole gearbox could be fuel-lubricated with no worries about leaks. The gas generator would run on LH2 the whole time due to its higher specific energy, but the chamber would ignite with methane alone at first and gradually shift the mix to hydrogen during the climb to orbit. 

[tater]

SSTOs have been proposed since the 60s. Not sure what the earliest was, but Phil Bono certainly put forth the SASSTO (Saturn Application SSTO) as a reusable SSTO. The thing was actually pretty tiny, with the payload as a Gemini capsule:

 

As in the OP, they used altitude-compensating—in this case a plug nozzle. I know a number of aerospikes have been tested, but as far as I know none have flown (firefly had been working on one, and abandoned it, right?)

Bono is interesting because in his book with Gatland, it's reuse, reuse, reuse. Clearly once operational reuse is a thing (turn around like an aircraft, or at least like a specialty aircraft—SR-71, U-2, etc), all bets are off on what might be cost effective. The SSTO component of a "rapidly reusable SSTO" (SSTOr as a shorthand vs an expendable SSTO?) seems substantially less important to me than the "rapidly reusable" part. If you have rapidly reusable TSTO vehicles, you could likely make an SSTO, though the margins for recovery from orbit are obviously going to be substantially tighter since it's going to have less mass to orbit to start with. That makes them possible, but more work to make it work, then optimize.

So I think the first operationally reusable vehicle will be TSTO—Starship, very likely. Once they demonstrate how to get back from orbit and turn around and do it again, then I think talking seriously about some sort of SSTO looks more sensible since the hard part—orbital recovery without so much wear that it needs to be substantially (in terms of labor, anyway) rebuilt—will have been sorted out.

We still have what payloads make sense for such a vehicle, and I still think it tends to be humans. 1000kg of humans is a decent number of humans (as long as you aren't selecting them at CostCo or Sam's Club ). Satellites? I think reuse is going to be a state change in what is possible in space, and heavier loads start looking more interesting—except for people. Robotics only improves, and the number of actual humans needed is not the same as 1950s and 1960s concepts. heck, not the same as O'Neill's ideas from the 70s. Heavy lift for cargo, smaller lift for humans. In that use case I think that maybe the reduced effort for a single vehicle might look attractive vs having to stack a TSTO.

If "rapidly reusable" is assumed to be an existing tech (maybe a few variant solutions are a thing), then the TSTO vs SSTO discussion has more interest to me, since we're comparing, well, costs. Someone makes an SSTOr, and someone makes a TSTOr, and what do they charge per kg? If one is substantially cheaper, does the more expensive one offer something that none the less makes it worthwhile that customers actually want?

 

Edited July 13, 2021 by tater

[DDE]

On 7/11/2021 at 12:51 AM, sevenperforce said:

There are a number of things you could do to simplify the design somewhat. Small performance hit but you could probably get better thrust out of it.

The RS-25 uses two fuel-rich preburners: one for the LOX pump turbine shaft and one for the fuel pump turbine shaft. Instead, you could just have a single preburner and split the exhaust out to two separate turbines, like the gas generator exhaust on the RS-68 and YF-77. Slight efficiency loss because the preburner exhaust would have a longer path to the turbines, but only using a single preburner would probably save weight anyway.

Or, hear me out.......

Expander cycle engines are simple and reliable, but they have poor thrust-to-weight and relatively low chamber pressure because there is only so much heat you can extract from the engine bell, and that gets worse and worse as your engine gets bigger.

However, what if you did a combination? Keep the same fuel-rich preburner for the LOX turbopump, but use a closed expander cycle for the fuel turbopump. Use the additional heat from the preburner to increase the thermodynamic potential in the expander cycle loop.

The LH2 would flow around the nozzle, chamber, AND preburner in order to absorb the maximum amount of heat. It would run through a turbine (which would turn the LH2 turbopump shaft), then exhaust into the preburner, which would burn with a small amount of LOX. The fuel-rich exhaust would run through another turbine (which would turn the LOX turbopump shaft) and then exhaust into the chamber.

The highest pressure would be in the LH2 coolant loop, which is an extremely gentle thermodynamic cycle in comparison to a preburner cycle. You still need a seal between the preburner turbine shaft and the LOX pump but that’s not nearly as challenging as dealing with an oxidizer-rich preburner. And since the preburner is only running the LOX turbopump, it is smaller and uses less energy than if you had one preburner for both pumps or two separate preburners.

Another approach, if you’re looking for a proper SSTO engine (for whatever reason), would be a variable-mixture tripropellant engine running LH2 *and* methane with LOX. The thermodynamic engine cycle for **that** would be insanity, though. Maybe you could do a hydrolox gas generator with exhaust to a LOX turbine and a fuel turbine, geared to both a LH2 pump and a methane pump. LH2 and methane play well enough together that the whole gearbox could be fuel-lubricated with no worries about leaks. The gas generator would run on LH2 the whole time due to its higher specific energy, but the chamber would ignite with methane alone at first and gradually shift the mix to hydrogen during the climb to orbit. 

Quote

Another design bureau, KB Khimautomatiki, worked on a third tripropellant engine based on their LOX/LH2 RD-0120 LPRE. The tripropellant concept was originally investi-gated and might have been originated by Rudi Beichel (originally of Peenemunde, Germany and later of Aerojet in the United States). The United States and other countries supported tripropellant engine studies and a few component tests, but never built the hardware for a complete tripropellant engine. The engine would be more complex, have probably six TPs (including booster turbopumps) and a more complex TC, be very expensive, and the vehi-cle performance improvement has been estimated to be small (maximum of 3%). The Russians have done some engine testing and were seeking interna-tional partners to share the cost of further engine development, estimated at more than $500 million. This is another example where the Soviets undertook three parallel developments for an application, which did not have a firmly designated future vehicle and which did not have a compelling mission advan-tage. None of these Soviet tripropellant engines has been fully developed. 

 

[Exoscientist]

On 7/7/2021 at 6:07 PM, sevenperforce said:

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?

 Thanks for the calculation.  As I math guy I love to see the calculations. We're probably pretty close in our estimates if you're taking, say, 350+ s for the Merlin vacuum Isp with "altitude adaptation"( below I'll give a link to a paper that uses this terminology.)

 The only difference is our interpretation of that 9,720 m/s delta-v you calculated.  For just going to LEO, that's actually pretty good. Such a vehicle could carry quite a bit of payload if you take the required delta-v to LEO as only 9,000 m/s. This is the value taken here:

Towards Reusable Launchers - A Widening Perspective.
https://www.esa.int/esapub/bulletin/bullet87/pfeffe87.htm

 Here's an article that examines various types of  "altitude adapting" nozzles:

Advanced Rocket Nozzles.
September 1998
Journal of Propulsion and Power 14(1998):620-633
Authors: 
Gerald Hagemann, Ariane Group 
Hans Immich, Aerospace Consultant
T. van Nguyen, Aerojet GenCorp, Inc.
Gennady Dumnov, Siemens

https://www.researchgate.net/publication/224796963_Advanced_Rocket_Nozzles

 The image shows how much higher an ideal altitude adapting nozzle can get in vacuum Isp compared to a standard bell nozzle. This is ideal but the point of the matter is rocket engines get quite high efficiency in the range of 95%+ 

 

  Robert Clark