SSTO Engine Technology

kbd512 · 2025-05-17 02:21:25

Dr Clark,

I decided to evaluate what an extendible / vacuum nozzle might net in terms of improved payload performance:

The Russian RD-0124 engine, a non-developmental LOX/RP1 engine in active service, has a vacuum Isp of 359s.

https://en.wikipedia.org/wiki/RD-0124

323.1s is 90% of that 359s Vacuum Isp

Mass Flow Rate (mdot) = Thrust / (Isp * g0)
3,452,113.5kg-f = 33,853,669N
mdot = 33,853,669N / (323.1 * 9.80665)
mdot = 33,853,669N / 3,168.528615
mdot = 10,684.35kg/s

6,508,946,390N-s / 33,853,669N = 192.267s
192.267s * 10,684.35kg/s = 2,054,248kg

Propellant Mass savings is 131,041kg, which nets an additional 7,407kg of payload to orbit. That amount of payload performance improvement would more than cover the mass allocation for the extendible nozzles. We need all the payload performance we can get for SSTOs, so I'll take it.

tahanson43206 · 2025-05-17 06:32:16

For kbd512 and RGClark ...

There must be a trade-off between the benefits of a variable geometry nozzle and just using two engines with the appropriate nozzles. All existing space access systems just use multiple engines instead of a variable geometry design.

It seems likely to me that the weight penalty of a second engine is pretty close to the weight any viable variable geometry scheme would have. GW Johnson has provided plenty (an abundance) of documentation on the challenges facing anyone who wants to try to create a variable geometry nozzle with known materials and techniques.

The second engine will need the plumbing to route fuel and oxidizer it's way, so that will definitely add mass.

Because it is possible readers of this post may not be familiar with GW's offerings, the bottom line is that the junction between the extension and the atmosphere bell is where a bit of tricky engineering is required.

GW tells me that if a vacuum engine is used for atmosphere propulsion, the tip will burn off right where the end of the atmosphere engine bell would be if the system were designed for atmosphere. If an extension is shifted forward to mate with the atmosphere bell, the junction is where the challenges will arise. GW seems to think it would be difficult to keep hot gases from burning through whatever seal your engineer team might come up with.

The hardware to shift the vacuum extension forward would have mass.

Why not simply design an SSTO with both engine types and eliminate the complexity of variable geometry?

The market opportunity I see is for a one person-to-LEO transport that is reusable, and a one way delivery of material comprising the vehicle itself to LEO for use in construction of a larger vessel such as the 500 passenger transport proposed by kbd512, or the 1000 person transport proposed by RobertDyck.

(th)

Last edited by tahanson43206 (2025-05-17 06:32:50)

GW Johnson · 2025-05-17 12:59:18

The trouble with mixing two engine types in an SSTO is the very same problem SpaceX's "Starship" would have if one tried to make it an SSTO: you are too short on thrust to begin with, limited by what will fit behind the stage. Adding a second engine type just makes that even worse! You have to have enough thrust to launch "smartly", which is takeoff thrust/weight near 1.5 or more (never less), or your gravity is loss increases dramatically. Why? You end up burning the vast majority of your propellant in only the first ~5 km off the pad, leaving you still moving very subsonic!

Before the recent design updates, "Starship" was in the neighborhood of 120 tons inert and 1200 tons propellant if fully filled. At ZERO payload, that's a launch weight of 1320 tons. It is normally configured with 6 Raptor engines, 3 sea level in the neighborhood of 250-300 tons sea level thrust (call it 275), and 3 vacuum engines which would be nearer 150-200 tons thrust (call it 175), if unseparated at sea level, which strongly limits exit bell expansion ratio.

By the way, raising vacuum expansion ratio with constant chamber and flow rate makes the exit area bigger, so that fewer engines like that, will actually fit behind the stage! The takeoff thrust problem really gets unsolvable very quickly, if you attempt what might otherwise seem obvious!

Thrust is mdot*Vexit + (Pe - Pa)Ae, where the pressure term is quite strongly negative at sea level for an unseparated vacuum engine. (Thrust would be almost zero with a separated bell, which would further destroy itself in a single handful of seconds.)

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

Last edited by GW Johnson (2025-05-17 13:29:10)

kbd512 · 2025-05-17 15:22:26

tahanson43206,

The RL-10B-2 engine uses a RCC extendible nozzle. This nozzle extension is about the same size and weight as what a Vacuum-optimized Merlin requires. It appears as though the real payload benefit to a hypothetical Space Shuttle mass SSTO vehicle so-equipped, is no more than about 2,300kg, based upon the nozzle hardware mass required to equip 40 Merlin engines. That's more than what I thought it would be, and most of the mass appears to be the nozzle extension itself, rather than the deployment mechanism, which has almost negligible mass per engine. The extension hardware is only 10 to 20lbs per engine.

Apart from the cost of the nozzle extension, the nozzle mass increase pretty much kills this idea. You would get more useful payload by cutting the weight of each engine by using the same RCC material for all the major engine components, combined with a staged combustion cycle. The Merlin is a marvel of gas generator engine tech, but staged combustion always provides higher Isp.

That 116,120kg mass to orbit value is inflexible because 6,508,946,390N-s delivers 116,120kg to orbit. We know this because that was the Total Impulse provided by 3X RS-25 engines affixed to the Space Shuttle and 2X SRBs.

When I use Silverbird Astronautics Launch Vehicle Performance Calculator, this is what I get:

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

When the dry mass is adjusted downward to only 13,426kg (4,315kg RCC engines + same 9,111kg propellant tank mass)
Estimated Payload: 91,813kg
95% Confidence Interval: 71,844kg - 116,164kg

ISS orbit, same RCC engines:
Destination Orbit: 400 x 400km, 45deg
Estimated Payload: 86,900kg
95% Confidence Interval: 67,807kg - 110,215kg

ISS orbit, same RCC engines, 323.1s Isp (90% of RD-0124 Isp):
Estimated Payload: 108,038kg
95% Confidence Interval: 86,343kg - 134,287kg

This appears to be little better or worse than the real Space Shuttle if we added the hardware for resuability. The only measurable performance improvement achieved during the STS program came from making the External Tank lighter. These payload performance estimates confirm that a modestly better Isp from the same engines and propellant mass confers a meaningful payload performance advantage, but only when greater dry mass doesn't immediately replaces that useful payload mass.

Historical Space Shuttle GLOW and Propellant Mass: 2,032,096kg; 735,602kg LOX/LH2 plus 997,904kg APCP, 1,733,506kg total
SSTO Space Shuttle GLOW and Propellant Mass: 2,301,409kg; 2,185,289kg LOX/RP1

The historical Space Shuttle has a propellant mass reduction of 451,783kg, but both LH2 and APCP are much more expensive than RP1.

The SSTO Space Shuttle carries 587,443kg of RP1. The real Space Shuttle carried 106,261kg of LH2 and 159,665kg of solid fuel within the APCP oxidizer / fuel combo mixed into the propellant grain of the solid rocket boosters, or 265,926kg in total. The remainder of the solid propellant mass was AP oxidizer.

LOX: $0.27/kg
RP1: $2.30/kg
APCP: $5.00/kg
LH2: $6.10/kg
LCH4: $8.80/kg
Hydrazine: $75.80/kg

SSTO Space Shuttle Fuel Cost
LOX: 1,597,846kg * $0.27/kg = $431,418
RP1: 587,443kg * $2.30/kg = $1,351,119
Total Propellant Cost: $1,782,537

Historical Space Shuttle Fuel Cost
LOX: 629,341kg * $0.27/kg = $169,922
LH2: 106,261kg * $6.10/kg = $648,192
APCP: 1,733,506kg * $5.00/kg = $8,667,530
Total Propellant Cost: $9,485,644

That makes the propellant costs 5.3X cheaper for the SSTO Space Shuttle vs the real Space Shuttle, despite the fact that the SSTO Space Shuttle is burning 452t of additional propellant. Most of what the SSTO variant is burning is LOX. RP1 exhaust doesn't produce HCl, either, unlike APCP. We lacked the materials and engine tech necessary for any kind of SSTO when the real Space Shuttle was designed, so that's a moot point.

If whatever changes you're making to the vehicle lead to a greater dry mass, then you need more engine power and more propellant, period.

kbd512 · 2025-05-17 15:25:51

GW,

Is there any reason why you cannot have multiple turbopumps feeding propellants into the same combustion chamber (so that you can use a much larger nozzle without singular gigantic turbopumps)?

GW Johnson · 2025-05-18 06:49:53

Kbd512:

There's no fundamental reason why you couldn't feed with from multiple turbopump assemblies. There might be a practical reason not to do that, though.

For the same flow velocities, flow rate should be proportional to cross section area, which suggests flow rate scales as pump dimension squared. Cutting down to half the flow in each pump would have those two assemblies each about 70% as large in dimension as the original single unit. That makes the packaging of the turbopumps about the engine chamber more crowded.

Since mass scales as dimension cubed, that option would be heavier, too.

I knew there was an extendible bell at least tested on some engine. From what you said, it was a variant of the RL-10. Did that ever fly on anything?

GW

Last edited by GW Johnson (2025-05-18 06:50:25)

RGClark · 2025-05-18 15:47:01

kbd512 wrote:

Dr Clark,
I decided to evaluate what an extendible / vacuum nozzle might net in terms of improved payload performance:
The Russian RD-0124 engine, a non-developmental LOX/RP1 engine in active service, has a vacuum Isp of 359s.
https://en.wikipedia.org/wiki/RD-0124
323.1s is 90% of that 359s Vacuum Isp
Mass Flow Rate (mdot) = Thrust / (Isp * g0)
3,452,113.5kg-f = 33,853,669N
mdot = 33,853,669N / (323.1 * 9.80665)
mdot = 33,853,669N / 3,168.528615
mdot = 10,684.35kg/s
6,508,946,390N-s / 33,853,669N = 192.267s
192.267s * 10,684.35kg/s = 2,054,248kg
Propellant Mass savings is 131,041kg, which nets an additional 7,407kg of payload to orbit. That amount of payload performance improvement would more than cover the mass allocation for the extendible nozzles. We need all the payload performance we can get for SSTOs, so I'll take it.

Yes, that would offer significant improvement. The problem is with a variable nozzle it’s not certain the “90% rule” would still provide an accurate estimate. You would need to do an accurate trajectory sim to be sure.

Bob Clark

RGClark · 2025-05-18 16:02:06

kbd512 wrote:

tahanson43206,
…
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
...

That mass ratio of nearly 100 to 1 may be too optimistic. The Falcon 9 first stage for instance gets about 20 to 1. You might be able to raise that to 30 to 1 using carbon-fiber tanks or the specialty high-strength steels on the Starship that SpaceX says matches carbon-fiber.

Bob Clark

kbd512 · 2025-05-18 17:40:30

GW Johnson wrote:

Kbd512:
I knew there was an extendible bell at least tested on some engine. From what you said, it was a variant of the RL-10. Did that ever fly on anything?
GW

This Day in Aviation - 5 December 2014, 12:05 UTC

From the article:

The Delta IV Heavy’s second stage is 42.8 feet (13.05 meters) long, and is also 16 feet, 10.0 inches in diameter. It uses an Aerojet Rocketdyne RL-10B-2 engine, producing 24,750 pounds of thrust (110.09 kilonewtons) of thrust. The RL-10B-2 is 13.6 feet (4.15 meters) long, 7.0 feet (2.13 meters) in diameter, and weighs 611 pounds (277 kilograms).

It looks like it flew to me, 11 years ago now:
Screen-Shot-2016-12-04-at-07.50.03.png.webp

It's going to fly again, or actually, "again again" (since it's already flown once on SLS as well), with the SLS ICPS stage for Artemis II.

Edit:
RL-10B-2

According to this sales brochure from Pratt & Whitney, the RL-10B-2 engine first flew in space 1999 aboard a Delta-III. That was about 26 years ago, and it was used by the Delta-IV Heavy as an upper stage engine. I'm going to assume this is very well established technology at this point.

Last edited by kbd512 (2025-05-18 17:49:49)

tahanson43206 · 2025-05-18 18:35:40

Post for link to cutaway of Rl-10b-2 engine (Extendible nozzle engine) (in use since 1999 Delta 3 rocket)

RGClark · 2025-06-23 22:42:29

GW Johnson wrote:

…
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

Can your software calculate the ISP vs. altitude of a sea level Merlin engine given adaptive nozzles? This graphic shows a radical improvement over the standard Vulcain using altitude compensation:

Russian kerosene upper stage engines have reached a max 360s vacuum ISP. So the sea level Merlin given an adaptive nozzles would increase its vacuum ISP from 312s to 360s or above.

Note such an ideally adaptive nozzles would also increase the sea level ISP, as the graphic shows for the Vulcain engine. The reason is fixed nozzle sea level engines are always overexpanded at sea level. This is because the engine designer also wants good performance in vacuum, so they select some intermediate expansion value. This reduces the sea level performance.

Note this means the adaptive nozzles also increases the sea level thrust over the standard engine. Then the adaptive nozzle has the twin benefit of increasing the vacuum ISP as well as reducing gravity drag due to increased thrust.

Bob Clark

Last edited by RGClark (2025-06-23 23:22:14)

kbd512 · 2025-06-24 09:45:32

Dr Clark,

How would you go about building an engine nozzle capable of changing its shape in 3 dimensions to match that 1D theoretically ideal expansion vs altitude graph, even now that we have lightweight materials able to withstand exhaust heat without regenerative cooling?

This is because the engine designer also wants good performance in vacuum, so they select some intermediate expansion value.

To the extent feasible, I would imagine that engine designers bias the nozzle expansion ratio toward the lower bound of the altitude range where the engine spends most of its firing time. If you spend far less time below 30,000ft than you do above 30,000ft, then it makes a lot more sense to optimize the nozzle's expansion ratio for 1/3rd of sea level pressure, however much you can get away with before significant flow separation occurs at sea level.

GW Johnson · 2025-06-24 13:55:50

Rocket nozzle thrust is F = CF At Pc, where Pc is understood to be the chamber pressure before flow starts necking down into the nozzle throat. Further, it is presumed that area contraction ratio is near 10, so that static pressure and total pressure at the Pc station are indistinguishable.

Here is the most convenient form of the CF equation: CF = (Ae/At)(Pe/Pc)(1 + g nKE Me^2) - (Pa/Pc)(Ae/At), in which Ae/At is the nozzle expansion ratio, Pa is the ambient backpressure, g is the specific heat ratio, nKE is the nozzle kinetic energy efficiency, and Pe is the expanded pressure in an unseparated nozzle, even if it is over-expanded.

There are two terms in CF: the first one (which is actually the vacuum CF) depends ONLY upon the expansion ratio Ae/At, which sets Me and Pe/Pc, for any given g, plus a value for nKE matching the detailed nozzle geometry. Me usually has to be determined iteratively from Ae/At. Once you know Me, you know everything: the compressible formulas are given in terms of Me and g, for Pe/Pc and for Ae/At.

The second term also depends upon Ae/At, but explicitly depends upon the actual values of Pc as well as Pa. This is the backpressure correction term for the retarding effects of being immersed in an atmosphere at some Pa. It is exactly zero in vacuum, because in vacuum Pa = 0 by definition.

The old standard sea level designs always used 1 atm for Pa, and set the expansion ratio Ae/At such that Pe = Pa. The backpressure term is significant: there is always less thrust at sea level, than out in vacuum, all else being equal. These Pe = Pa sea level nozzles have all the thrust you are going to get at sea level, but only somewhat more in vacuum, because the Pe = Pa design limits your expansion ratio rather strongly. It is the Pe = Pa design approach that makes Isp sensitive to Pc, via that backpressure term.

You can use any of a variety of larger expansion ratios at sea level, as long as you do not separate the bell due to excessive Pa compared to Pe. When you do that, the thrust you get at sea level reduces below the old-time sea level design’s thrust, but the vacuum thrust you get with it is larger, because the expansion ratio is larger. That increases the average ascent Isp of the nozzle, flying from sea level out into vacuum, but the cost for that is the reduced thrust at sea level, just at the time when you are the very heaviest at launch, trying to fly straight upward against gravity!

The dilemma is that you are always short of thrust at launch, limited by how many engines of a given thrust will actually fit behind the stage. I cannot emphasize that enough! It is why just “doing the rocket equation” will ALWAYS lead you astray when you do actual design sizing work.

The limit for separation is determined by any of a number of empirical (!!!!) correlations over the years. The one I like is very simple to use, and determines Psep/Pc = (1.5*(Pe/Pc))^0.8333. At any given design with a specific Pc value, you then know the value of Psep. If Pa > Psep, the bell separates (and is usually destroyed in a matter of a few to only several seconds). As long as Pa < Psep, there is no separation, although the backpressure term at sea level does strongly reduce your CF because of the larger Ae/At, and thus your thrust.

The trend in recent years toward higher Pc values is easily explained looking at the thrust equation F = CF At Pc. At the same expansion ratio and CF, higher Pc lets you use a smaller At for the same thrust F! Smaller At is smaller Ae and smaller engine length. More of these higher-Pc engines will fit behind the stage, getting you more thrust at sea level launch when you need it the most.

And, yes, Bob, you can use my spreadsheet to model that extendible-bell RL-10 variant. Run the sea level design and determine the flow rate and throat area. Then re-run the problem with the bigger expansion as a vacuum design, and adjust your vacuum thrust sizing value until you get exactly the same flow rate and throat area as the sea level design. Pick an altitude to do the switch-over. Use the sea level's performance data vs altitude up to that point, then use the vacuum's performance data vs altitude from that point on up to vacuum. Then combine those data into one plot. Do the averaging of Isp across altitude with the same two sets of data, but combined into one set at the switch-over altitude, and that’s a pretty good approximation to the ascent-averaged Isp you would see.

No, it is complicated, you simply don't do this with a single or simple calculation. It takes lots of calculations, which is exactly what the spreadsheet does for you. But you can do it.

GW

GW Johnson · 2025-06-24 16:23:12

I took a look around at several sites regarding the RL-10B-2 with the extendible exit cone. I did find it was only used as an upper stage vacuum engine. The stowed nozzle allows it to fit in a shorter interstage space, but it is not fired that way! The ONLY data for it are in the extended vacuum configuration. That's not to say a 2-piece extendible nozzle could not be designed and developed, but the RL-10B-2 is NOT one of those!

Most applications have more interstage length available, which is why the other variants, including the newer RL-10C's, have fixed geometry vacuum nozzles. It's only used for upper stage service as a vacuum engine these days. The low-expansion RL-10A's are all long-retired.

GW

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#26 2025-05-17 02:21:25

Re: SSTO Engine Technology

#27 2025-05-17 06:32:16

Re: SSTO Engine Technology

#28 2025-05-17 12:59:18

Re: SSTO Engine Technology

#29 2025-05-17 15:22:26

Re: SSTO Engine Technology

#30 2025-05-17 15:25:51

Re: SSTO Engine Technology

#31 2025-05-18 06:49:53

Re: SSTO Engine Technology

#32 2025-05-18 15:47:01

Re: SSTO Engine Technology

#33 2025-05-18 16:02:06

Re: SSTO Engine Technology

#34 2025-05-18 17:40:30

Re: SSTO Engine Technology

#35 2025-05-18 18:35:40

Re: SSTO Engine Technology

#36 2025-06-23 22:42:29

Re: SSTO Engine Technology

#37 2025-06-24 09:45:32

Re: SSTO Engine Technology

#38 2025-06-24 13:55:50

Re: SSTO Engine Technology

#39 2025-06-24 16:23:12

Re: SSTO Engine Technology

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