Copyright 2023 Robert Clark
Abstract.
Most orbital rockets have payload fractions in the range of 3% to 4%. The Ariane 6 using 2 and 4 SRB’s, because of the large size of the SRB’s and because solids are so inefficient on both mass ratio and ISP, the two key components of the rocket equation, it will count among the worst rockets in history at a payload fraction of only 2%.In contrast a two Vulcain Ariane 6 could have a payload fraction of 7% and a three Vulcain Ariane 6 could have a payload fraction of 7.5%. This is well-above what any other rocket has ever achieved in the history of space flight.
So how is an all-liquid Ariane 6 able to accomplish this? First, this version is based on the Ariane 5 core. The mass ratio for the Ariane 5 it turns out is quite extraordinary for a hydrogen+liquid oxygen(called “hydrolox”) stage at 16.3 to 1. This is in the range commonly seen by dense propellants. To use a colorful analogy, it’s like the ArianeSpace engineers in designing the Ariane 5 core found a way to make liquid hydrogen as dense as kerosene!
Obviously, this is not what happened. But they must have found a way to achieve extreme lightweighting of a hydrolox stage. To put this in perspective, the mass ratio of the famous Centaur hydrolox upper stage is at 10 to 1, achieved back in the 1960’s. And the Delta IV hydrolox core is at a quite ordinary 8.7 to 1 mass ratio. So the Ariane 5 core is about twice as good as the Delta IV core on this key mass ratio scale.
Because the Ariane 5 core has the high Isp of a hydrolox stage while achieving (somehow!) the high mass ratio of a dense propellant stage, it calculates out to have the highest delta-v of any rocket stage in the history of spaceflight.
Since delta-v is the single most important parameter for orbital rockets, you can legitimately argue the Ariane 5 core is the greatest rocket stage ever produced in the history of spaceflight.
The high 7.5% payload fraction of the all-liquid Ariane 6 would mean SpaceX would have to be chasing ArianeSpace rather than the other way around.
To put this advance in perspective, it would be like SpaceX using the very same Merlin engines and the very same propellant tanks, and the very same size Falcon 9, suddenly being able to change the Falcon 9 payload from 22 tons to 40 tons.
It will represent a paradigm shift in terms of the payloads that rockets will be expected to deliver to orbit.
Usually, when we think of a radical shift in rocket capability we imagine some great advance in engines such as nuclear, or some great advance in materials to greatly reduce tank weight.
Quite extraordinary is the the fact this radical increase in rocket capability can come from using currently existing engines and tanks.
Mitchell Burnside Clapp is an engineer and former Air Force officer who had been prominent in programs back-in-the-day to find low cost space access, such as the DC-X. Here, he was calculating some existent or previous space stages that just on the ideal delta-v parameter would have SSTO capability:
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Propellant density, scale, and lightweight structure
.
lightest aerospace structures ever built. The thing that keeps
it from being a single stage to orbit machine is its relatively
heavy and low performance engines.
I decided to examine the historical record on this issue and
developed the table you see below. All the weights are in
thousands of pounds; the Ideal DV colum is in kft/sec. Prop wt
refers to the weight of propellants. The Isp numbers are
referenced as far as possible to vacuum Isp. The engine data are
from CPIA Revised Liquid Propellant Engine Manual (1972), and
Rocketdyne, SEP, Aerojet, and Pratt and Whitney product
information sheets. The vehicle weight data are from Isakowitz’
Space Launch Systems, 1st Ed. Ideal DV is calculated from the
rocket equation [DV = Isp * g * ln (gross/(gross-prop))]. The
column labeled PMSMF refers to the Propulsion-Free Structural
Mass Fraction, which refers to the weight of the stage after the
engine is removed, divided by the gross weight. This residual
weight includes the electronics, tankage, and so forth.
Here are the data:
Stage Prop Wt Gross Wt Engine Wt Isp PFSMF Ideal DV
Titan II Stg 1 260.0 269.0 3.258 287 2.13% 31.372
Black Arrow Stg 1 28.7 31.1 1.426 250 2.99% 20.755
Saturn V Stg 1 4584.0 4872.0 93.080 265 4.00% 24.114
Titan III Stg 1 294.0 310.0 3.343 283 4.08% 26.959
Titan IV Stg 1 340.0 359.0 3.343 283 4.36% 26.731
Delta 6925 Stg 1 211.3 223.8 2.528 295 4.46% 27.383
Atlas E 248.8 266.7 4.371 312 5.07% 27.073
Saturn V Stg 2 993.0 1071.0 17.400 425 5.66% 35.821
Zenit Stg 1 703.0 778.0 26.575 337 6.22% 25.364
Titan III Stg 2 77.2 83.6 1.144 312 6.29% 25.796
Saturn IB Stg 2 233.0 255.0 3.480 425 7.26% 33.504
Titan II Stg 2 59.0 65.0 1.102 308 7.54% 23.611
Saturn IB Stg 1 889.0 980.0 16.072 263 7.65% 20.111
Ariane 5 Stg 1 342.0 375.0 3.630 430 7.83% 33.624
Saturn 5 Stg 3 238.0 263.0 3.480 425 8.18% 32.179
Energia Core 1810.0 1995.0 21.000 452 8.22% 34.583
Zenit Stg 2 178.0 198.0 2.480 350 8.85% 25.816
Black Arrow Stg 2 6.5 7.8 .531 265 9.52% 15.450
Titan IV Stg 2 77.2 87.0 1.144 312 9.95% 21.919
Delta 6925 Stg 2 13.4 15.4 .207 267 11.82% 17.416
...
7. There are several stages that have SSTO-class delta-V figures
(anything over 30000 fps). The Titan II first stage can itself
deliver 1400 pounds to low earth orbit as it sits, with no
modifications to engine or structure. That’s pretty impressive,
even if a load of propellant for it costs $2.5 miilion.
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https://groups.google.com/g/sci.space.policy/c/PZgWB9WWhNw/m/gWAavQL8AAAJ
However, since the SSTO was dismissed as not worth-while an important implication of such high delta-v stages was missed: when the first stage gets such high delta-v, the upper stage can be much smaller to get the same payload to orbit.
Or said another way, a high delta-v first stage gets high payload to orbit with just a small size upper stage.
Of those high delta-v stages Burnside Clapp listed as of 1995, the Ariane 5 is the only one yet existent. And in actuality its even much better than listed in the table.
You see, the Ariane 5 core had a forward skirt called the JAVE("Jupe AVant Equipée") that transmitted the thrust of the two side boosters to the core. Without the side boosters, this would be removed in our version. The JAVE weighed 1,700 kg. So lets calculate again the ideal delta-v without the JAVE. Note I'm using the lighter Ariane 5 "G" stage here, rather than the later "E" version, at a 158 ton propellant load and 12 ton dry mass. Removing the JAVE brings down the dry mass to 10.3 tons. I'll use the slightly better 434s vacuum Isp for the Vulcain now rather than the 430s Mitchell Burnside Clapp used in his 1995 calculation. Then the ideal delta-v is:
434*9.81Ln(1 + 158/10.3) = 11,900 m/s, or 39,000 ft/sec.
This is by far the best ideal delta-v ever produced by any single rocket stage in the entire history of spaceflight, exceeding also the delta-v's for the separate rocket stages on the Falcon 9.
How were the Europeans able to produce such an extraordinary rocket stage? Firstly, they used hydrogen/oxygen(hydrolox) propellant on the core; this is known to produce the highest efficiency on the ISP scale of any chemical propellant. But this is well known and you see several stages among the highest listed used hydrolox. What's really extraordinary is the mass ratio, i.e., propellant fraction of the stage.
The mass ratio is gross mass divided by dry mass and it is important for a rocket stage for it is used in the rocket equation to determine what is the delta-v it could achieve:
The mass ratio of the Ariane 5 core for a hydrogen stage is so remarkable that it should be used as a model for any stage hydrolox or kerolox.
Cost of the side SRB's is the source of the high Ariane 6 pricing.
That parameter of ideal delta-v that the Ariane 5 core has superiority on over any other rocket ever built suggests that that should be built upon and not disregarded. Instead what has been used on the Ariane 5 and Ariane 6 are solid stages that are among the worst on this key parameter. Because solids are pressure-fed meaning the entire propellant tank has to operate as the combustion chamber requiring thick tank walls, they usually have poor mass ratio, despite the solids propellants higher density. Worse, the other key parameter in the rocket equation ISP is also among the worst with solids.
If it were only small solid side boosters used with the Ariane 5 and Ariane 6 then this would not have the severe reducing effect on the rocket efficiency that they did have. But instead the side rockets used on the Ariane 5 and 6 were huge in comparison to other side boosters used for example with the Atlas 5 and Delta IV, which were commonly only ~1/10th the size of the booster core stage.
My speculation here, but I think the Space Shuttle design is what influenced ArianeSpace to use such large solid side boosters. Perhaps it was not known at the time when the Ariane 5 was first being designed but the Space Shuttle was a financial disaster. Such large side boosters are also used on the SLS which is also a financial disaster. The huge solid boosters used in both contributed to that.
The newly designed solids on the Ariane 6 make the situation worse. Their size is about the size of the entire core stage of the Ariane 6, and the fact they use carbon-fiber make them more expensive. To understand how expensive is that use of carbon-fiber for the solids note the reason SpaceX decided to move away from carbon-fiber to steel for the StarShip:
Why SpaceX Abandoned Carbon Fiber.
…
The other concern was cost. SpaceX determined that it would spend upwards of $130,000 per ton to use carbon fiber as the primary rocket body material. On the other hand, it would spend just $2,500 per ton for stainless steel. It doesn’t take a mathematician to figure out that spending 50 times as much on carbon fiber would put considerable strain on the Starship project.
https://markets.rockwestcomposites.com/we-now-know-why-spacex-abandoned-carbon-fiber
To provide an estimate of how bad is the cost issue against the Ariane 6 solids in comparison to just using an additional Vulcain, note the €75 million cost of the two SRB version of the Ariane 6 compared to the €115 million of the four SRB version. Then, as a first order estimate, we can take the cost of two SRB’s as €40 million. But the cost of a single Vulcan is only €10 million! So the two SRB’s planned for the base version costs 4 times as much as just adding a second Vulcain!
Therefore, again as a first order estimate, we can take the cost of a Ariane 6 with no SRB’s by subtracting off the estimated €40 million for the two SRB’s to get a no SRB price of only €35 million.Then the price of the two SRB's is more than the price of the entire rest of the rocket. So adding on a Vulcain at €10 million would give a price of €45 million, about $50 million. Note this compares quite favorably with the current $67 million cost of the Falcon 9 new.
Further indication of how expensive are the Ariane 6 SRB's is found by comparing to other carbon-fiber, also called graphite-fiber, SRB's. The GEM 63 are carbon-fiber solid side boosters have about a 50 ton propellant load and cost estimated in the range $5 to $7 million.Then we can estimate the Ariane 6 SRB's to cost three times more to bring them to $15 to $21 million each, in the price range of the estimate you get from comparing the Ariane 6 two SRB and Ariane 6 four SRB pricing.
Payload Calculation for a Two Vulcain Ariane 6.
In the blog post, "Multi-Vulcain Ariane 6", I estimated about 11 tons to LEO using a two Vulcain, no SRB version for the Ariane 6. Note though I was originally trying to find a lower cost approach to the Ariane 6, so I actually used the Ariane 5 core. BUT because of my thrust constraints I chose to use the original, somewhat smaller version the Ariane 5 "G" core, rather than the later "E" version, at 12 ton dry mass and 158 propellant mass.
For the upper stage, again because of limited take off thrust constraints I did not use the current ESC-A cryogenic upper stage of the Ariane 5 at ~19 ton gross mass, nor the ~30 ton cryogenic upper stage of the Ariane 6. I also did not like that the ESC-A had such a poor mass ratio at only 5 to 1. I used instead the Ariane 4's H10 cryogenic upper stage:ARIANE 4 STAGE 3
Specifications are given in H10/H10+/H10-3 order.
Designation: H10/H10+/H10-3
Engine: single cryogenic open cycle SEP HM-7B
Length: 10.73 m/11.05 m/11.05 m
Diameter: 2.60 m
Dry mass: 1,200 kg/1,240 kg/1,240 kg, excluding interstage 2/3
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 10,800 kg/11,140 kg/11,860 kg
Thrust: 63 kN vac/63.2 kN vac/64.8 kN vac
http://www.braeunig.us/space/specs/ariane.htm
Note that in addition to being lighter this has a much better mass ratio at over 10 to 1, rivaling the famous Centaur upper stage. I also assumed the Vulcain thrust could be ramped up ca. 9% as was shown possible with the SSME's and the RS-68 engine on the Delta IV rocket. Then I asserted for the Vulcain likely the same would hold, as they are all hydrolox engines.
Launcher analysis and cost benefits
But even without a higher thrust level Vulcain, we could take a smaller propellant load of ca. 140 tons for the Ariane 6 core so it could still have sufficient thrust for takeoff. Since this is only about a 10% reduced propellant load for the first stage it would be a relatively small reduction in payload.
While I used the smaller, earlier Ariane 4 H10 upper stage because of my reduced thrust, we can now use the higher efficiency and thrust Vinci engine rather than the original HM-7B on this upper stage. The Vinci has a 180 kiloNewton vacuum thrust and 457s vacuum Isp. But using a slightly longer nozzle we can give it the 465.5 vacuum Isp of the RL10-B2. Remarkably with a sufficiently long nozzle we can give an upper stage hydrolox engine a vacuum Isp in the 480+s range. However, we'll use in our calculations the Isp number proven possible with currently in use engines of a 465.5s vacuum Isp.
As for the dry mass of the core with a second Vulcain, we removed the JAVE subtracting off 1,700 kg from the dry mass. Adding on a second Vulcain adds on 1,800 kg, bringing the dry mass back to about the original 12 tons.
However, another consideration is the doubled thrust might require thickened tank walls. I estimated in an earlier blog post that the increased thrust might require an additional 1,000 kg for the thicker tank walls. However, it should be noted the supported weight that needs to be carried above the core with a smaller upper stage and smaller payload mass is half as big as that of the Ariane 5. So advanced structural analysis programs need to be applied to find the needed degree of tank strengthening and added weight to the dry mass.
Still at most 1,000 kg needs to be added to the dry mass according to my prior estimate and this results in a proportionally small reduction in the payload mass. Then for this first order estimate we'll take simply 12,000 kg as the dry mass.
Now use the payload estimator at SilverbirdAstronautics.com giving the results:
I have presented the two Vulcain case here and in some previous postings because of the low development cost, less than $200 million. Indeed, it most likely could be done for less than $100 million.
However, adding two additional Vulcains will require a higher development cost. It still likely will be in the few hundred million dollars range, well less than the multi-billion dollar development cost of the current version of the Ariane 6.
Again a key consideration is added tank wall thickness needed for the tripled thrust. In a prior blog post, I discussed a rocket that had been proposed by Northrup Grumman, the Liberty rocket that would use a shuttle derived SRB as a first stage, a la the Ares I rocket, and an Ariane 5 core as an upper stage.
As described in this video, the SRB to be used would have had a thrust 12 times that of the Vulcain yet the increased thickness of the tanks on the Ariane 5 core would only need to be 50%:
As the tank mass of the Ariane 5 core is in the range of 4 tons, this would mean the increased tank mass would have needed to be in the range of 2 tons. Since the three Vulcain format would mean far less less thrust than that of the Liberty rocket, the increased tank mass would be less than this. So we'll take the additional core mass as a max of 2,000 beyond that of the additional 3,600 kg for the added two Vulcains.
Since we have higher thrust we'll use the larger Ariane 5 "E" core at 170 propellant load and 14 dry mass. Subtracting off again the 1,700 kg for the JAVE, while adding on 3,600 for the two added Vulcains and 2,000 kg for the thickened tank walls brings the dry mass to about 18,000 kg. But because we have much more liftoff thrust with three Vulcains we can use much larger upper stages, such as the currently planned 30 ton hydrolox upper stage of the Ariane 6, or even larger 40 ton or 50 ton hydrolox stages.
We'll take our upper stage as 50 tons propellant load with a Centaur-like 10 to 1 mass ratio, so a 5 ton dry mass, a la the ULA Centaur V. We'll use three Vinci engines on the upper stage to give a thrust of 540 kN and assume a RL10-like 465.5s Isp. Then the input page on the SilverbirdAstonautics.com payload estimator appears as:
.JPG)
And the GTO payload is:
.JPG)
On this key parameter rocket engineers use to rate orbital rockets the all-liquid Arianes would literally be the best rockets in the history of space flight with no other rocket even coming close.
They would literally be a paradigm shift in rocket efficiency. Other launch companies would have to strive to reach their level of efficiency. And most simply could not.
To put this advance in perspective, it would be like SpaceX using the very same Merlin engine and the very same propellant tanks, and the very same size Falcon 9, suddenly being able to change the Falcon 9 payload from 22 tons to 40 tons.
Plus, while being nearly twice as good as the Falcon 9 on this key parameter the all-liquid Ariane 6 would also be cheaper!
Robert Clark
