Stoke Space

[darthgently]

49 minutes ago, tater said:

Except an aerospike masses more, and takes up more room. So in return for a few seconds of I_sp gain, it loses TWR, the trade is not worth it. Firefly started with a toroidal aerospike... and abandoned it.

The only use-case for an aerospike I think is for an SSTO.

But what does every single rocket company on earth know?

Agree.  Maybe with further advances in rotating detonation the eqn will change, but hard to beat light simplicity over heavier complexity

What about a vertical version of an aircraft carrier steam launch system to get that first 5 to 10 m/s in "stage 0" bwauahahaha

[magnemoe]

1 hour ago, tater said:

Except an aerospike masses more, and takes up more room. So in return for a few seconds of I_sp gain, it loses TWR, the trade is not worth it. Firefly started with a toroidal aerospike... and abandoned it.

The only use-case for an aerospike I think is for an SSTO.

But what does every single rocket company on earth know?

Who is an excellent point. for an reusable first stage you want high TWR at launch as you want as you want to get the second stage up to its velocity fast, this doubles down if you want to return to the launch site as you then need to do the boost back burn who you want to do as fast as possible. Add that the 3.5 meter diameter is nice for road portability so I say the next diameter is 6 meter. 

 

47 minutes ago, darthgently said:

Agree.  Maybe with further advances in rotating detonation the eqn will change, but hard to beat light simplicity over heavier complexity

What about a vertical version of an aircraft carrier steam launch system to get that first 5 to 10 m/s in "stage 0" bwauahahaha

Rotating detonation will probably be more relevant for upper stages, they burn for longer and don't need serious trust and you are in vacuum who make things easier. 

[tater]

The reality is that for reusable rockets—particularly those designed for rapid reuse ("rapid" here just means little or no refurb between flights)—you are not optimizing for payload mass fraction, you are optimizing for COST. The primary drivers on reusable vehicles are fixed operational costs, which you at some point have no control over, they reach whatever the min is, and propellant cost.

If the props are sufficiently cheap, who cares if you burn a little more fuel to get to LEO? Buying a few % efficiency while making other things more expensive makes no sense. The same holds for people obsessed with SSTOs. An SSTO buys you nothing if it costs vastly more than TSTO (unless it has some benefit worth cost, like safety, rapid reflight, etc).

[Exoscientist]

20 hours ago, tater said:

I understand trying to come up with better variants of crappy rockets that suffer from design-by-politics (Ariane 6, SLS, etc), but I don't understand second guessing people who have clearly thought about all the trade offs for a vehicle like the one Stoke is building.

Andy Lapsa explains why an aerospike is a poor booster choice in the video posted up thread (the aerospike discussion is timestamped, but I linked at that timestamp):

 

 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

Edited June 20, 2023 by Exoscientist
Clarity

[tater]

1 hour ago, Exoscientist said:

 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.

Linear aerospike. How many of those are you going to put on the bottom of a ~3.7m rocket?

Stoke's rocket is smaller than F9 payload wise ("much more" than 1.6t, so we don't know, and of course in comparing to F9 we'd need to actually count S2 mass since reused), but how much thrust do they need? XRS-2200 is roughly comparable to 1 Merlin. But it's 3.4m long, by 2.3m wide. So 1 of them replaces 2 Merlins. Not sure how you combine linear aerospike engines, seems to me it works on Venture Star which is WIDE, but you'd not want them like V V, so no putting multiples on the bottom. So this gets you a rocket that must mass ~57t gross on the pad.

Actually, even for annular designs, can you cluster them? I'd imagine not.

If so, you are stuck with a single, large aerospike, or a WIDE launcher with a linear aerospike. The RS-2200 (full size for Venture Star) was to be over 9m long.

Edited June 20, 2023 by tater

[sevenperforce]

On 6/19/2023 at 10:10 AM, Exoscientist said:

On 6/14/2023 at 11:42 AM, tater said:

The booster is only giving the thing a couple km/s. Yes, they'd save a few percent of booster props from the average I_sp being higher.  It doesn't buy enough. Vacuum I_sp for methalox is an upper limit something like 380s. Sea level is ~330s. A SL methalox engine running in vacuum gets maybe 350s. So the difference is ~30s of I_sp, and the average will be somewhere in the middle.

I'd be mystified if they did an aerospike.

The XRS-2200 engine showed you can get quite high vacuum ISP of even a first stage engine by using an aero spike.

The XRS-2200 engine was built from the J-2S powerplant, which is not a first stage engine at all. The J-2 was developed from the ground up as high-energy upper-stage engine to replace the cluster of eight RL-10A-3S engines on the Saturn I S-IV stage. Because the RL-10's closed expander cycle cannot be readily adapted to produce a reliable engine with greater than 155 kN, Rocketdyne designed the J-2 to use a gas generator design to achieve around 53 bar chamber pressure and produce over 1 MN of thrust.  The expansion ratio wasn't great -- just 27.5:1 and no nozzle extension -- but the 53 bar chamber pressure was respectable.  It was never, ever a first stage engine. Firing it at sea level would have achieved only 200 seconds of specific impulse.

By using a more simplified cycle with the same expansion ratio, the J-2S reached 15 seconds greater vacuum specific impulse than the J-2 and eliminated about 84 pounds of unnecessary mass, increasing the vacuum TWR from the J-2's 73.2:1 to just over 85:1. It was still never a first stage engine. The J-2X, an updated design with a proper vacuum nozzle extension, would have achieved 448 seconds of specific impulse.

The XRS-2200, on the other hand, went back to the gas generator cycle of the J-2 but split the propellant flow up between ten separate combustion chambers. Smaller combustion chambers and a bigger nozzle meant that the XRS-2200 could achieve similar 436.5 seconds of specific impulse, similar to the J-2S but still far short of what could be expected from the J-2X's true vacuum nozzle. The aerospike nozzle meant that it could be fired at sea level without serious losses or flow instability and still achieve 339 seconds of specific impulse. That's something, but it's still much lower than the RS-68A's 363 seconds of sea level specific impulse. The biggest problem with the XRS-2200? It packed on the pounds. At 3,450 kg, it was more than double the mass of the J-2 despite only producing 15% higher vacuum thrust. The added weight came both from the small clustered nozzles, which suffered from the square-cube law, and the extremely heavy aerospike ramp nozzle. The sea level thrust to weight ratio would have been just 21.2:1, barely half of a true sea level gas generator hydrogen engine like the RS-68A.

On 6/19/2023 at 10:10 AM, Exoscientist said:

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.

The first stage of a two stage vehicle gains thrust due to underexpansion as it climbs into vacuum. Again, using the RS-68A as an example, the specific impulse at liftoff is 363 seconds while the specific impulse in a vacuum is 412 seconds. Assuming a similar launch profile to the Delta IV Heavy, a first stage powered by the XRS-2200 would take a full minute to catch up to the specific impulse of the RS-68A. By that time the booster has burned a full quarter of its propellant. And you'll need twice as much engine to achieve the same TWR when you're using the XRS-2200 because aerospikes are so heavy, which means you either accommodate increased gravity drag or you accommodate increased dry mass.

You can't just say assume that it will increase the overall performance; you actually have to look at a real-world example.

On 6/19/2023 at 10:59 AM, Exoscientist said:

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.

For the record, anything past about 20 km is essentially vacuum for the purposes of nozzle optimization.

But again, an aerospike engine is NOT a vacuum-optimized engine. An aerospike gets lower vacuum specific impulse than a vacuum-optimized engine and lower sea level specific impulse than a sea-level optimized engine, and it will be significantly heavier than either. If you want to make a proper comparison, you'll need to integrate over the entire first-stage burn sequence.

On 6/19/2023 at 5:05 PM, darthgently said:

On 6/19/2023 at 10:59 AM, Exoscientist said:

See this graphic of a Delta IV flight for example.

The first stage starts at 0m/s and ends somewhere much nearer to, but short of, orbital velocity, so the actual time spent at lower altitudes compared to higher altitudes is much greater than the diagram you posted suggests, as it is simply the path taken and the horizonal axis is not time

Actually working up the numbers is challenging. You have to pick a launch TWR and you have to use numeric integration. The thing you want to minimize is total Δv loss, where total Δv loss is the sum of gravity drag, aerodynamic drag, and specific impulse shortfall (also known as pressure drag), added to the Tsiolkovsky Δv shortfall resulting from dry mass differences.

You can make it easier of course if you pick a known-optimized trajectory and choose the same TWR.

On 6/20/2023 at 10:14 AM, Exoscientist said:

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.

An adaptive nozzle on a first stage that would give it the same vacuum Isp as a vacuum optimized nozzle and the same sea level Isp as a sea level optimized nozzle would improve the overall rocket performance if the Tsiolkovsky Δv losses were low enough.

That doesn't exist. There is no engine, aerospike or otherwise, with an adaptive nozzle that achieves the same sea level specific impulse as a sea level nozzle as well as the same vacuum specific impulse as a vacuum nozzle. And even if it DID exist, you would have to also factor in dry mass increase with the Tsiolkovsky rocket equation.

[sevenperforce]

On 6/20/2023 at 2:37 PM, Exoscientist said:

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.

As I pointed out above, the SSME (the RS-25) achieves high vacuum specific impulse because it is a vacuum-optimized engine. It has an adaptive nozzle which allows it to fire safely at sea level, but at the expense of significant thrust losses.

Anyhow, you're also comparing apples and oranges. First, you're talking about staged combustion specifically, not closed cycles generally (closed expander cycles, for example, tend to have very low chamber pressures). Second, "closed cycle" and "high chamber pressure" are not linearly related. The closed-cycle RS-25 boasts 206 bar, almost double the chamber pressure of the gas generator RS-68, but the closed-cycle BE-4 is only 17 bar higher than the gas generator Vulcain 2. Finally, cost does not directly correlate.

On 6/20/2023 at 2:37 PM, Exoscientist said:

Both of these are in the range of a $10 million cost for example

Where are you getting that number? In 2006, the open-cycle RS-68s were $20 million each -- about $31 million in today's dollars. And it doesn't correlate. The closed-cycle BE-4 is $8 million, far cheaper today than the RS-68 was twenty years ago. 

On 6/20/2023 at 2:37 PM, Exoscientist said:

And with the closed cycle, high chamber pressure SpaceX is using on the Raptor, they still haven’t gotten it to operate properly. Based on the number of engines that failed prior to the launch of the Superheavy/Starship about 1/3rd of the Raptors fail.

I'm really lost as to what this has to do with vacuum specific impulse, but regardless, intentionally testing a design to failure is not how you come up with failure rates. 

On 6/20/2023 at 2:37 PM, Exoscientist said:

In contrast SpaceX was able to get the low pressure, open cycle Merlin to operate reliable in fairly short order.

Merlin's 108 bars of chamber pressure would like very much to know who you are calling low pressure, and politely directs your attention to the closed-cycle YF-75 and RL-10A's ~42 bar. 

On 6/20/2023 at 2:37 PM, Exoscientist said:

Stokes approach to use several small thrusters on the upper stage should be the same approach they use on the lower stage.

They already appear to be using a cluster of engines on the lower stage, so I'm not sure about your point here.

On 6/20/2023 at 2:37 PM, Exoscientist said:

 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. 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.

The engines on their second stage are far, far too small to be used on an appropriately-sized first stage.

On 6/20/2023 at 2:37 PM, Exoscientist said:

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.

The XRS-2200 occupied a forward area of 7.8 square meters, giving it a sea level thrust/area ratio of 117 kN/m². The RS-68A has a sea level thrust/area ratio of 676 kN/m². So no, it doesn't work, not for this application. It doesn't work at all. And that's before we even start talking about the horrendous dry mass problems.

[tater]

22 minutes ago, sevenperforce said:

They already appear to be using a cluster of engines on the lower stage, so I'm not sure about your point here.

Not only that, they're gonna try full-flow staged-combustion.

[sevenperforce]

38 minutes ago, tater said:

1 hour ago, sevenperforce said:

They already appear to be using a cluster of engines on the lower stage, so I'm not sure about your point here.

Not only that, they're gonna try full-flow staged-combustion.

Which @Exoscientist thinks is a poor choice because of cost, complexity, and schedule.

The issue is that for optimizing any two-stage launch vehicle, the first stage liftoff TWR is a more significant factor than the first stage vacuum specific impulse. Playing funky games with nozzles can help with vacuum specific impulse, but it doesn't improve liftoff TWR. Increasing chamber pressures will improve liftoff TWR while also giving better vacuum specific impulse as a bonus.

Edited June 21, 2023 by sevenperforce

[tater]

35 minutes ago, sevenperforce said:

The issue is that for optimizing any two-stage launch vehicle, the first stage liftoff TWR is a more significant factor than the first stage vacuum specific impulse. Playing funky games with nozzles can help with vacuum specific impulse, but it doesn't improve liftoff TWR. Increasing chamber pressures will improve liftoff TWR while also giving better vacuum specific impulse as a bonus.

Yeah, that particular choice is certainly interesting, but they are trying to be more of a "fast follower" I assume—least on the booster tech. Raptor is an existence proof for them.

Stage 2 is pretty next level, can't wait to see it fly.

[sevenperforce]

6 minutes ago, tater said:

Yeah, that particular choice is certainly interesting, but they are trying to be more of a "fast follower" I assume—least on the booster tech. Raptor is an existence proof for them.

Stage 2 is pretty next level, can't wait to see it fly.

Stage 2 really seems cool.

I think there could be a way to do a partially aerospike-based first stage, maybe using some sort of multi-cycle architecture or tripropellant trickery.  But unless it was for an SSTO or sustainer, I can't see that it would be worth the trouble. 

[Exoscientist]

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

 

[sevenperforce]

13 hours ago, Exoscientist said:

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.

If you look closely at the graphic above, you'll see that the Stoke engine is an expander bleed engine, not a closed expander like the RL-10. While it functions on the same principle, it's a fundamentally different engine cycle in terms of performance.

As a rule, you shouldn't expect two engines to have similar chamber pressures merely because they use the same propellant combo and fill the same role -- not unless they are also the same engine cycle. The Soyuz-FG's RD-107 and the Soyuz-5's RD-171 are both first-stage engines running on kerolox, but the first has a chamber pressure of 61 bar while the second has a chamber pressure of 250 bar. The YF-75D on the second stage of China's LM5 uses hydrolox and gets 41 bar, while the YF-90 on the second stage of its LM9 also uses hydrolox but gets 183 bar. The AJ-10 is an upper-stage hypergolic engine with a chamber pressure of 8.6 bar while the Vikas is an upper-stage hypergolic engine with a chamber pressure of over 50 bar. Differences in engine cycle matter much more than whether the engine is plopped onto the first stage or onto the second stage.

13 hours ago, Exoscientist said:

A 100 bar chamber pressure can serve as a legitimate, fixed nozzle first stage engine, i.e., no aerospike required.

A 61 bar chamber pressure can serve as a legitimate first stage engine, as noted above with the RD-107.

13 hours ago, Exoscientist said:

the RS-68 on the Delta IV . . . gets 412s vacuum ISP. And the Vulcain on the Ariane 5 . . . 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.

Specifically, that's because the Vulcain 2 has an expansion ratio of 60, almost triple the expansion ratio of the RS-68A. In this, it is very much like the RS-25; it's essentially a vacuum nozzle shortened just enough to fire at sea level. This demonstrates the inherent tradeoff in engine design, and emphasizes why an aerospike can't simultaneously function as a perfectly efficient sea level engine and a perfectly efficient vacuum engine.

13 hours ago, Exoscientist said:

without aerospike Stoke Space can use the same upper stage thrusters on the first stage

If it wants to use a vastly large first stage, sure. Hydrogen is fluffy. But it wants reuse, and reusing a first stage requires optimization for high thrust, so that you can get into space and get back quickly before you are too far downrange.

13 hours ago, Exoscientist said:

To get their “foot in the door” I think Stoke Space should try initially. . . .

Stoke Space doesn't want to "get their foot in the door"; Stoke Space wants to solve operational reuse of both stages.

13 hours ago, Exoscientist said:

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.

Simpler and cheaper if you don't care about second-stage reuse, yeah.

13 hours ago, Exoscientist said:

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.

The first stage has to lift the second stage, too. And deal with gravity drag. Which is why the first stage will need 6-10 times as much thrust (and thus 6-10 times as many engines of equivalent thrust) as the second stage.

13 hours ago, Exoscientist said:

Stoke could reduce the thrusters on the upper stage from 15 to, say, 10. . . .

. . . scrapping any possibility of their main goal which is operational second-stage reuse . . .

13 hours ago, Exoscientist said:

. . . while using the originally planned 30 thrusters . . . on the first stage. . . .

. . . which would never get off the ground, because they would need 60+ thrusters to provide the appropriate balance, and closer to 80+ thrusters if they want enough oomph to get off the ground quickly for first-stage reuse.

[tater]

15 minutes ago, sevenperforce said:

Stoke Space doesn't want to "get their foot in the door"; Stoke Space wants to solve operational reuse of both stages.

QFT

First words on their webpage:

Quote

We’re on a mission.

It starts with 100% reusable rockets.

 

[Exoscientist]

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

 

[SunlitZelkova]

15 hours ago, Exoscientist said:

I advise stick with the small launcher first before proceeding to the large launcher.

Why? Is there any good reason beyond an arbitrary imposition of another company’s methods on this one?

[magnemoe]

16 hours ago, Exoscientist said:

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

Hydrogen engines has superb ISP but low trust.  Most hydrolox rockets uses SRB. Is it any other rockets than Delta 4 heavy who was pure hydrogen? 
And their engine does not look that advanced. I say their main benefit is size. 
They can land second stage anyplace who allow it and just have an truck pick it up and take it to repairs for next trip.  You don't do that with an starship as its massive.  KSC is probably the only place they might land it outside an offshore platform? 
 

[Exoscientist]

10 hours ago, SunlitZelkova said:

Why? Is there any good reason beyond an arbitrary imposition of another company’s methods on this one?

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

  Robert Clark

[magnemoe]

36 minutes ago, Exoscientist said:

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

  Robert Clark

Who is correct, they started with the V2 rocket.  And for many space companies it madesprefect sense to start with an small rocket, cheaper to develop and cheaper to operate. Launching small satellites was also an lucrative business. Today that marked is pretty cramped. I also assume its plenty of engineers who has experience with larger rockets  that is makes little sense starting with an smaller rocket if it don't make profit. 
Now starting with an reusable second stage on the other hand is ambitious. 

[SunlitZelkova]

15 hours ago, Exoscientist said:

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

  Robert Clark

So do you think SpaceX should have started with sounding rockets? Because that is what the governmental launch programs actually started with.

[darthgently]

16 minutes ago, SunlitZelkova said:

So do you think SpaceX should have started with sounding rockets? Because that is what the governmental launch programs actually started with.

https://www.google.com/search?tbm=isch&q=goddard&tbs=imgo:1#imgrc=Jet1rH2SeM5UGM

[tater]

41 minutes ago, SunlitZelkova said:

So do you think SpaceX should have started with sounding rockets? Because that is what the governmental launch programs actually started with.

Not to mention Blue Origin. Look how they have moved on from that start over their 23 years of existence!

[Exoscientist]

4 hours ago, tater said:

Not to mention Blue Origin. Look how they have moved on from that start over their 23 years of existence!

 Even Blue Origin started with small suborbital New Shepard.

  Robert  Clark

[darthgently]

2 hours ago, Exoscientist said:

 Even Blue Origin started with small suborbital New Shepard.

  Robert  Clark

But have they "moved on" much in 23 years?

[Exoscientist]

 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