Exoscientist

I’m going by Lapsa’s statement in the video where he literally says this stage can go to 400,000 feet, which is suborbital space. So he must mean taking into account the fact the thrust and Isp levels of the engines are reduced at sea level at takeoff. There are several European start-ups on the horizon planning to take advantage of the small launch market of < 1,000 kg to LEO. With the miniaturization of satellites, this market is expected to be sizable. With it’s level of development Stoke Space could probably beat these European start-ups to market joining Rocket Lab as the only companys offering such launches.(The ESA’s Vega could also but it’s price of $37 million is prohibitive.) By the way, Stoke really wasn’t so innovative with it’s cone shaped stage nor with its plug nozzle for cooling on reentry. This was thought up back in the 60’s by noted rocket designer Philip Bono and even patented, though the patent has probably lapsed by now: ToughSF @ToughSf A history of VTOL SSTOs, the 'platonic ideal' of space launch vehicles: https://researchgate.net/publication/253467497_History_of_the_Phoenix_VTOL_SSTO_and_recent_developments_in_single-stage_launch_systems… Key features are a plug-nozzle aerospike engine, integrated heatshield and very lightweight structures+tanks. 4:00 PM · Dec 26, 2022 · 8,137 Views 16 Retweets 1 Quote 124 Likes 17 Bookmarks https://twitter.com/toughsf/status/1607481507869851652?s=61 The examples shown there including the famous DC-X show a ground-launch conical stage is not prohibitive against a stage reaching suborbital space or even being a SSTO. Quite notable as well is Bono wanted to use the efficiency of the aeroplug/aerospike for the launch, not disparage it. Bob Clark

Below are the input page I used and results page for the Silverbirdastronautics.com calculator. For the Vulcain engine, I took the vacuum thrust as 1,350 kN. So for three Vulcains I input 4050 kN in the thrust field for the first stage. I took the vacuum ISP as 434 s. Since it had higher thrust than two Vulcains I used the later, larger version Ariane 5 “E” core at 170 ton propellant load and 14 ton dry mass. I added ~4 tons for two additional Vulcains and another ~2 tons for strengthening the tank for the higher thrust, but also subtracted ~2 tons for removing the JAVE. The resulting dry mass I used was ~ 18 tons for the first stage. For the second stage, the higher thrust enabled a larger upper stage. So I took it comparable to the Centaur V at a 50 ton propellant load and 5 ton dry mass. For the engines I used three Vinci engines at 180 kN vacuum thrust each for a total of 540 kN, but I mentioned I assumed a larger nozzle to reach the highest 465.5 s Isp of the RL10. The result was 19.6 tons to LEO. This is nearly 3 times your result of 7 tons. May I see the input page you used to the Silverbirdastronautics.com since the inputs you used must have been very different from the ones I used? Robert Clark

ATK, now Northup Grumman, partnering with Astrium proposed a modification of the Ares 1 that would use a Shuttle derived SRB for the first stage and an Ariane 5 core for the upper stage called the Liberty rocket. The Astrium engineers would remove the JAVE attachment at the top of the core without the SRB’s of the Ariane 5: Liberty. A single Vulcain 2 engine powers the stage, providing 136 tonnes of thrust for about 540 seconds. It also provides roll control during the main propulsion phase. At shut down, EPC separates from Ariane's upper stage and reenters. On Ariane 5, the EPC is topped by a forward skirt named JAVE ("Jupe AVant Equipée") that transmits thrust from the two solid boosters to the core vehicle. This structure will not be needed on Liberty. Araine 5 launchers are topped by a Vehicle Equipment Bay (VEB) that houses guidance and control systems, along with sets of hydrazine fueled roll and pitch thrusters. The VEB, which is usually positioned atop the Ariane 5 second stage, is 5.4 meters in diameter, is 1.56 meters tall, and weighs 1.9 tonnes. EPC will be modified for use on Liberty. The tank walls will have to be strengthened, but some or all of the added mass would likely be offset by the elimination of the JAVE. Vulcain 2 would have to perform an air start, but since it is a gas generator engine modifications are expected to be managable. Snecma and Astrium have already been working on air start designs. Ground test firings of an air-start Vulcain 2 would occur in mid 2013 if Liberty won a NASA contract. https://web.archive.org/web/20160301144316/http://www.spacelaunchreport.com/liberty.html I agree with you the ballon tank structure of the Ariane 5 core was a key reason why it was able to achieve high propellant mass ratio. But remember the famous Centaur upper stage also uses ballon tank structure and has been in use over 50 years, with very high reliability. About the payload estimates using the SilverbirdAstronautics.com estimator, as described in its documentation you have to use the vacuum values of both the thrust and ISP even for the first stage engines. The reason is the calculator already takes into account the diminution at sea level. Robert Clark

As I mentioned in the blog post, on the Ariane 5 core the JAVE on the forward skirt of the Ariane 5 core is specifically to transmit the high thrust of the two side boosters: https://www.esa.int/Enabling_Support/Space_Transportation/Launch_vehicles/Cryogenic_main_stage_EPC This is unnecessary without the two side boosters. Removing this 1,700 kg is how I get the 16.3 to 1 mass ratio. I encourage you to do the ideal delta-v calculation of this Ariane 5 core with a 434s vacuum Isp and 16.3 mass ratio. It is higher than any other stage ever existing including the Saturn V upper stages and including either stage of the Falcon 9. Using this high mass ratio Ariane 5 core, calculate the estimated payload to LEO by the rocket eq. of a two Vulcain, no-SRB version of the Ariane 6 using the lighter Ariane 4 H10 upper stage at ~10 ton prop. load to enable lift-off with the lower thrust without SRB’s. Then for a three Vulcain, no-SRB version, use a larger Centaur V like upper stage enabled by the higher take-off thrust and estimate the payload by the rocket eq. in that case. The payload fraction is worsened with the SRB’s because their mass is so high you get a high number in the denominator, reducing the fraction even if the payloads are similar. The low mass of the hydrolox stages in contrast puts a smaller number in the denominator, so increases the fraction size. Robert Clark

At about the 30 minute mark in the Everyday Astronaut video Andy Lapsa mentioned the upper stage they are going to do their hops test with can actually by the rocket equation make 400, 000 feet, 120 km, which is past the von Karman line for suborbital space: They decided against the suborbital space test because they would have to change the “outer mould line”, presumably changing the tapered cone shape. Perhaps for launch of this stage to reach the needed Mach speed for suborbital space, a straight cylindrical shape would be optimal for reasons of the effects of supersonic drag. But remember how SpaceX swapped out the rings on the Starship on a regular basis? This really isn’t that major a modification. I don’t know though if this capability of suborbital space flight is with the 15 engines or the original 30 engine complement. Remember for the hop tests you can just leave it partially fueled so it can take off to low altitude hops with fewer engines. I’ll assume it can reach the suborbital space altitude with the full 30 engines. From this released image of the upper stage tank we can estimate its dimensions by comparing to the height of the men in the photo at approx. 6 feet tall: I estimate the bottom width of the tank as 11.4 feet, and the other dimensions proportionally. Taking the density of hydrolox as approx. 300 kg/m^3, I estimate the propellant mass as approx. 10 tons. SpaceX used 9 Merlins on the Falcon 9 booster and 1 on the second. Other commercial start-ups for small launchers such as Rocket Lab and Relativity Space are following this pattern, 9 engines on the booster and 1 on the second stage. Then follow this pattern for Stoke Space with that completed upper stage now acting as the booster. With this booster stage having 30 engines, take a new smaller upper stage as having 3 to 4 engines. Commonly the upper stage of a TSTO is 1/3rd to 1/4th the size of the first stage. So take the size of the new upper stage as 2 to 3 tons. The famous Centaur hydrolox upper stage at a 20 ton propellant load got an approx. 10 to 1 mass ratio. But that doesn’t necessarily mean a smaller stage can get that same weight efficiency. In fact, scaling a rocket stage up usually improves mass ratio. So logically scaling it down should in general reduce it. Still the weight efficiency of some the ArianeSpace stages has been impressive. I discussed on the ArianeSpace thread the Ariane 5 core at a 158 ton propellant load got a surprising hydrolox mass ratio of 16.3 to 1. And the Ariane 4 H10 hydrolox upper stage at only ~10 ton propellant load got an approx. 10 to 1 mass ratio. So it is possible for a 10 ton hydrolox stage to get an approx. 10 to 1 mass ratio. So I’ll take Stoke’s now booster as 10 ton propellant mass and 1 ton dry mass. I’ll assume the ground launch engines can match the 434 s vacuum Isp of the Vulcain on the Ariane 5. But what about the upper stage? I’ll take the propellant load as 3 tons. But again a scaled down stage won’t necessarily have the same high mass ratio. Instead of a 10 to 1 mass ratio, I’ll take it for this smaller stage as approx. 6 to 1, with a 0.5 ton dry mass. For it’s vacuum Isp I’ll take it as the RL10’s best 465.5 s. Then I can get approx. 1.0 ton, 1,000 kg, to LEO by the rocket eq.: 434*9.81Ln(1 + 10/(1 + 3.5 + 1.0)) + 465.5*9.81Ln(1 + 3/(0.5 + 1.0)) = 9,400 m/s, probably sufficient for LEO. Note SpaceX and Rocket Lab offered the Falcon 1 and the Electron at a few hundred kg to LEO capability at the $8 million to $12 million range. And the new commercial launch start-ups in Europe expect to offer their small 1 ton to LEO launchers also in this price range. It seems likely Stoke Space could match or beat this price point. Robert Clark

NASA suggests Artemis lander flight likely pushed back to 2026 due to Starship delay. The NASA official quoted suggests needed launches for qualifying Starship plus the needed launches for the refueling flights will likely delay the Artemis landing flight: NASA predicts delay: Starship grounded pending investigation By Steve Clark - June 28, 2023 https://myrgv.com/local-news/2023/06/28/nasa-predicts-delay-starship-grounded-pending-investigation/ Robert Clark

NASA suggests Artemis lander flight likely pushed back to 2026 due to Starship delay. The NASA official quoted suggests needed launches for qualifying Starship plus the needed launches for the refueling flights will likely delay the Artemis landing flight: NASA predicts delay: Starship grounded pending investigation By Steve Clark - June 28, 2023 https://myrgv.com/local-news/2023/06/28/nasa-predicts-delay-starship-grounded-pending-investigation/ Bob Clark

Strictly speaking, it was the Everyday Astronaut in his video on the Stoke rocket that suggested the chamber pressure would be 100 bar. Andy Lapsa himself never said that. What Lapsa did say was that the upper stage engines would be “low pressure”. See at about the 28 minute point in the video: It is debatable if a rocket engineer such as Lapsa would consider 100 bar to be “low pressure”. Because it is an upper stage engine running hydrolox I’m inclined to think he meant “low pressure” a la the RL10 at ca. 40 bar. I think most rocket engineers speaking in general terms would call 100 bar a mid-level engine. On the other hand by “low pressure” he may have meant in comparison to their planned full flow staged combustion engine for the first stage, likely to be > 200 bar chamber pressure. Anyway, as far as their current plans they’re not going to use these engines on the first stage on their operation rocket anyway, only for test hops. The speculation about using an aerospike or aero plug on the first stage is only relevant if they did use these upper stage engines on an operational first stage. And even then it’s utility would be doubtful if the chamber pressure for these is in fact 100 bar, as Everyday Astronaut suggests, since you can get fairly high vacuum ISP with a hydrolox engine at only 100 bar as the Ariane 5 Vulcain demonstrates. Robert Clark

I’m not so sure about that for two reasons. First, the core has to withstand the full thrust of the engine even without the SRB’s when they are jettisoned. Second, the core has a structure called the JAVE("Jupe AVant Equipée") that transmits the forces of the two side boosters to the core. That structure weighed 1,700 kg. So that structure would be unneeded without the SRB’s so the core could be made even lighter. That lighter weight without the JAVE is what I used in my calculation. By the way, I’ve been informed by someone familiar with the Ariane 5 construction that the core tank weighed ca. 4,400 kg. But it was made of aluminum. But you can save about 50% off tank weight over aluminum using carbon fiber or the specialty high strength stainless steels SpaceX is using on the Starship. So you could subtract an additional 2,200 kg from the core stage mass by using these lightweight materials for the tank. That would bring the dry mass down to 8.1 tons. But that would mean the mass ratio of the core would be over 20 to 1. Robert Clark

Actually, that would be useful as an initial test of the tech. Robert Clark

Yes. The discussion at Stoke was over suborbital space. The decision was against it to only low altitude hops. I advise proceeding to test flights to suborbital space after the short hops. Then the very same stage that managed the suborbital space flights could serve as the first stage of a two stage to orbit launcher. Robert Clark

Actually, that might not be a bad approach for Stoke. The first three launches of the Falcon 1 failed. And Elon Musk has admitted if the fourth had failed then they would have had to close shop since they would have been out of money. It is notable that Rocket Lab first started with suborbital flights. Once they succeeded at that, they proceeded to the orbital case. It’s not a coincidence that their first three orbital launches of the Electron were successful.(*) Stoke had speculated at a high altitude test of the Hopper, but decided on low altitude tests. After succeeding with these low altitude tests, I’d advise proceeding to high altitude, essentially to suborbital space. Thereafter proceed to the two-stage orbital case. Indeed the stage that did the flights to suborbital space could serve as the first stage of the all-hydrogen two stage rocket to orbit. (*)Note: the first launch of Electron was proceeding perfectly according to the data returned by the rocket. However, the U.S. Air Force negotiated with Rocket Lab and with New Zealand that they would have “minders” during the launch, with the authority to send a destruct signal if they thought necessary. It turned out the Air Force software for the radar track of the rocket was flawed and lost sight of the rocket. They commanded to send the destruct signal even though the rocket was actually following the correct track. This is described in the nice book: When the Heavens went on Sale. https://www.amazon.com/When-Heavens-Went-Sale-Geniuses/dp/B0BCD4D4DK/ref=tmm_aud_swatch_0?_encoding=UTF8&qid=1687808044&sr=1-1 Robert Clark

No. The point is they did a suborbital first. Blue Origins sluggish approach to development is decidedly odd. It can be argued their slow approach contributed to ULA’s financial difficulties because the Vulcan Centaur was so badly delayed by the delay in the development of the BE-4 engine. Robert Clark

Blue Origin motto is, “Graditum Ferociter!” It translates as “Step-by-step ferociously!” It’s been noted by industry observers Blue is focused far too much on the “Graditum” and almost none on the “Ferociter”. Robert Clark

Even Blue Origin started with small suborbital New Shepard. Robert Clark

EVERY orbital launch program started with small launchers first including the billion dollar governmental launch programs of the U.S. and Russia. Robert Clark

SpaceX had to build the Falcon 1 before it built the Falcon 9. And Rocket Lab had to build the Electron before it could proceed to the Neutron. The initial plan for Stoke also was for a 1.6 ton payload as fully reusable, though a twitter post from Stoke suggest a higher number now. I advise stick with the small launcher first before proceeding to the large launcher. Rocket Lab also like SpaceX is using both the same propellant and same engines other than nozzle size on both stages of their rockets. This offers both simplicity and lower cost, obviously two very big considerations for a start-up. This article suggests the full flow staged combustion may be too ambitious for their first engine Stoke builds for the first stage: Stoke Space to build SpaceX Raptor engine’s first real competitor. https://www.teslarati.com/stoke-space-spacex-starship-raptor-engine-competition/ SpaceX has already spent billions on the SuperHeavy/Starship and still hasn’t gotten the Raptor to operate reliably. The key fact is by using the same engines Stoke already has for both stages Stoke essentially can be flying test missions now. In contrast it is extremely unlikely Stoke will be able to get a full flow staged combustion engine to operate reliably before SpaceX with billions of dollars at their disposal can. Robert Clark

In the Everyday Astronaut video on Stoke Space a graphic shows the upper stage engines having a combustion chamber pressure of 100 bar. That surprised me, since being hydrolox upper stage engines I expected them to be of similar chamber pressure as the RL10, at ca. 40 bar. A 100 bar chamber pressure can serve as a legitimate, fixed nozzle first stage engine, i.e., no aerospike required. For instance the RS-68 on the Delta IV first stage has an approx. 100 bar chamber and gets 412s vacuum ISP. And the Vulcain on the Ariane 5 first stage also at approx. 100 bar pressure gets 434s vacuum ISP. The RS-68 needs better thrust at sea level, in not having the large side boosters of the Ariane 5, so sacrifices vacuum ISP, getting more sea level thrust. So without aerospike Stoke Space can use the same upper stage thrusters on the first stage. To get their “foot in the door” I think Stoke Space should try initially an all-hydrolox two stage system. Remember SpaceX opted for using the same fuel and the same engine, aside from nozzle size, for both stages on the Falcon 9. This was the simpler and cheaper approach. Stoke following this approach, means the difference between the upper and lower stages, aside from size, would be just the number of thrusters used. Commonly the first stage is three to four times the size of the second stage. So Stoke could reduce the thrusters on the upper stage from 15 to, say, 10, while using the originally planned 30 thrusters of the upper stage to instead 30 thrusters being used on the first stage, and a proportionally larger first stage than the second. Robert Clark

I added the following abstract to the blog post: Towards a revolutionary advance in spaceflight: an all-liquid Ariane 6. 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, which was 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 say 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. https://exoscientist.blogspot.com/2023/06/towards-revolutionary-advance-in.html Robert Clark

He makes two arguments there; one is correct the other incorrect. The first point that you can get a high vacuum ISP from a ground launch engine by using a high chamber pressure is correct. This is for example done with the SSME. But the SSME using a closed cycle is an expensive engine. Open cycle engines such as the RS-68 on the Delta IV or the Vulcain on the Ariane are much simpler and cheaper. Both of these are in the range of a $10 million cost for example, while the SSME is of the range of $50 million. (Aerojet Rocketdyne in “redesigning” them for the SLS absurdly made them even more expensive at $125 million each.) And with the closed cycle, high chamber pressure SpaceX is using on the Raptor, they still haven’t gotten it to operate properly. Yes I know they get them to fire, but a rocket engine leaking fuel and catching fire is NOT normal, certainly not for an operational engine. Based on the number of engines that failed prior to the launch of the Superheavy/Starship about 1/3rd of the Raptors fail. And by failing, keep in mind for a good number of them that means actually leaking fuel and catching fire. SpaceX claimed the Raptor 2 used on the SH/SS test launch was more reliable. In the test launch 1/4th of them failed, including at least two that actually exploded. That’s not going to cut it. In contrast SpaceX was able to get the low pressure, open cycle Merlin to operate reliable in fairly short order. Furthermore as SpaceX showed its much cheaper to use several small rockets on a rocket stage then a single large one. So Stokes approach to use several small thrusters on the upper stage should be the same approach they use on the lower stage. Then for cost reasons and speed to operational status they should use multiple low pressure, open cycle engines just as they are doing on the upper stage. But then the argument of about getting the high vacuum Isp by using a high pressure closed cycle engine no longer applies. Note, also this means the engines they already have can be used also on the first stage, rather than waiting for an expensive high pressure, staged cycle engine for the first stage. So that’s the first point he mentioned. But the second point he discusses is incorrect. That’s the one about a too small throat area for the aero spike. The reason it doesn’t apply is because that is for a large single aerospike engine. But that is not the case being discussed here. The case being discussed here is the multiple low pressure, open cycle case. And we know that case does work. How do we know it works? Because it was already built and tested 20 years ago by NASA in the XRS-2200. Bob Clark

As I said even the Falcon 9 first stage fires into quite high altitude, ca. 80 km. It is the case for any two-stage vehicle the first stage fires into near vacuum altitude, where the vacuum ISP is the one that would obtain. Then an adaptive nozzle on a first stage that would give it the same vacuum Isp as a vacuum optimized nozzle of an upper stage would improve the overall rocket performance. Robert Clark

Towards a revolutionary advance in spaceflight: an all-liquid Ariane 6. http://exoscientist.blogspot.com/2023/06/towards-revolutionary-advance-in.html Most 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. 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. It will be a paradigm shift in what payloads rockets should be able to deliver to orbit. Robert Clark

Actually a key fact is a significant proportion of a first stage firing is under near vacuum conditions, where a vacuum optimized engine would improve its performance. See this graphic of a Delta IV flight for example. The first stage fires all the way to 120 km altitude. The Falcon 9 first stage also fires until quite high altitude, nearly at vacuum, at ca. 80 km. Bob Clark

They will first test the upper stage on the ground, a “hopper”. The upper stage will use hydrogen so these first hopper experiments will use hydrogen. The XRS-2200 engine showed you can get quite high vacuum ISP of even a first stage engine by using an aero spike. Then quite key is to recognize a large portion of the flight of even a first stage is under near vacuum conditions so an aero spike can increase performance of a first stage of a two stage vehicle, thereby increasing the performance of the rocket overall. Robert Clark

Who got the engines? Bob Clark