Stoke Space

[tater]

8 hours ago, Exoscientist said:

 Even Blue Origin started with small suborbital New Shepard.

So if Stoke takes more than 23 years (24 years, assuming NG flies next year, else 25 years) to get to orbit, they will have demonstrated your point.

 

[Exoscientist]

5 hours ago, tater said:

So if Stoke takes more than 23 years (24 years, assuming NG flies next year, else 25 years) to get to orbit, they will have demonstrated your point.

 

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

Edited June 26, 2023 by Exoscientist

[Exoscientist]

18 hours 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.

 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

 

Edited June 26, 2023 by Exoscientist

[tater]

Stoke IS starting with suborbital.

Stage 2 will undergo hopping tests.

If you define "suborbital" to require achieving some arbitrary altitude, then no.

They have zero reason to make some small, useless rocket. Smallsat launch is a dead end.

[magnemoe]

1 hour ago, tater said:

Stoke IS starting with suborbital.

Stage 2 will undergo hopping tests.

If you define "suborbital" to require achieving some arbitrary altitude, then no.

They have zero reason to make some small, useless rocket. Smallsat launch is a dead end.

I disagree that Smallsat launches is dead, its just an marked so many fight for, some might launching at loss just to get an foothold. 

And suborbital in my book is getting close to 80 km. Or I just did an manned suborbital mission jumping up 20 cm  
Or more seriously nobody claims mortars or 155 mm artillery is suborbital, nor battleship guns, rocket artillery with 500 km range and you don't look like an idiot saying if its a ballistic missile.   

And as I assume Stoke is not much above an small sat launcher anyway.  That upper stage is pretty small and fueled with hydrogen. 

[tater]

55 minutes ago, magnemoe said:

And as I assume Stoke is not much above an small sat launcher anyway.  That upper stage is pretty small and fueled with hydrogen. 

Someone said it was 1.6 tons to LEO, and the response from them was way more than 1.6t. Not defining what "way more" meant. I think of a few hundred kg as small sats.

Functionally, their first vehicle is in fact a small rocket. Stage 2 reuse will be easier with a larger vehicle for a number of reasons. Once they get some launches and EDL under their belts, they can scale up.

[Exoscientist]

5 hours ago, magnemoe said:

I disagree that Smallsat launches is dead, its just an marked so many fight for, some might launching at loss just to get an foothold. 

And suborbital in my book is getting close to 80 km. Or I just did an manned suborbital mission jumping up 20 cm  
Or more seriously nobody claims mortars or 155 mm artillery is suborbital, nor battleship guns, rocket artillery with 500 km range and you don't look like an idiot saying if its a ballistic missile.   

And as I assume Stoke is not much above an small sat launcher anyway.  That upper stage is pretty small and fueled with hydrogen. 

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

[tater]

The only reason to go any higher than low altitude hops is if at can hit reentry velocities to test the heat shield. It's a waste of time.

[Exoscientist]

7 hours ago, tater said:

The only reason to go any higher than low altitude hops is if at can hit reentry velocities to test the heat shield. It's a waste of time.

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

  Robert Clark

[tater]

2 hours ago, Exoscientist said:

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

No, it wouldn't, as stage 2 is never going anywhere by itself. That's what stage 1 is for.

[magnemoe]

2 hours ago, tater said:

No, it wouldn't, as stage 2 is never going anywhere by itself. That's what stage 1 is for.

Agree, first stage engines just have to have enough TWR to get it into orbit and enough to land it empty. This limit how much fuel you can launch it with as you need TWR>1. 
But its enough for your to do jump for practicing landings. 
For reentry testing, you are now in orbit, so deploying an customers satellite, you can now try to reenter and land the thing. 
SpaceX is not even try to recover the first reentry attempts for Starship. Stoke second stage is small enough that you can land it at some remote place like the military testing areas and send an truck to pick it up.

[tater]

We know what Stoke's current plan is, they have told us. They don't have to guess at the mass of their vehicle, or anything else that is required to successfully test it—they are designing it to fly, and in what limited data we have they seem to be pretty thoughtful about their path. Second-guessing them seems presumptuous to me.

They were founded in 2019.

If this was 2029, 2039, or 2042 and they had not yet done anything aside from some suborbital hops—yeah, I'd start second-guessing them.

[tater]

 

[Exoscientist]

On 6/21/2023 at 3:18 PM, sevenperforce said:

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.

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. 

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. 

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. 

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

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

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.

 

  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

[sevenperforce]

5 hours ago, Exoscientist said:

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:

<snip>

It is debatable if a rocket engineer such as Lapsa would consider 100 bar to be “low pressure”.

If Lapsa never said it was 100 bar, then we have no idea what the chamber pressure is, although I am guessing Tim isn't just making that number up from nothing. However, the comment on low pressure for the upper stage engine is a commentary on engine cycle, not engine role, because he is talking about the structural characteristics of the thrust chambers.

Having all-copper thrust chambers isn't JUST due to pressure; even the RL-10 has a thrust chamber that is structurally steel, and we don't imagine the Stoke upper stage engine is lower-thrust than the RL-10. (The entire thrust chamber and upper nozzle of the RL10 is made of stainless steel tubes that are painstakingly brazed together; I believe it has an internal copper jacket but I could be wrong.) The key here is to understand that this is NOT a typical or conventional engine, so direct comparisons aren't appropriate. Recall that one of the major issues with multichamber aerospikes is the square-cube law: as an object shrinks, the volume shrinks faster than the area, making it heavier for its size. However, one (small) saving grace is that as an object shrinks, you get a mechanical advantage in terms of necessary tensile yield strength.

Stoke can tolerate all-copper combustion chambers because the design pressure is lower than their larger staged-combustion engine design pressure AND because the individual chambers themselves are so physically small. It's a bad idea to extrapolate endlessly from a one-off comment.

5 hours ago, Exoscientist said:

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.

As I noted above, there is absolutely no reason to think that rocket engineers segregate between upper-stage and lower-stage engines in considering what is low pressure and what is high pressure; they segregate between engine cycles. Engine cycle is primary. A 55 bar closed expander would be a high pressure closed expander, but a 55 bar gas generator would be a low pressure gas generator. A 130 bar bleed expander (like the MB-XX) is a high pressure bleed expander, but a 140 bar staged combustion engine would be a low pressure staged combustion engine.

5 hours ago, Exoscientist said:

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.

Or just on the lower end for open cycle engines generally. He does, after all, follow this up IMMEDIATELY by talking about engine cycle: "It helps the open cycle be more efficient, because you don't need as much energy to drive the turbines." In an open cycle like a bleed expander (or, really, any open cycle engine, including gas generators), you can drive chamber pressures arbitrarily high by dumping as much propellant as you want overboard through the turbine: the more propellant you dump, the more energy you'll have at the power head. But of course dumping prop means losing specific impulse. And while you gain some specific impulse with a greater chamber pressure, given a fixed nozzle size, you'll quickly reach a point where the prop loss outweighs the increased specific impulse (not to mention the greater turbine mass and heavier thrust chambers you'll need).

This doesn't mean Tim's diagram is correct and the chamber pressure is 100 bar. It might be lower. It might be higher. It certainly doesn't add any meat to speculation about what would happen if they were a completely different company building a completely different rocket for a completely different purpose.

5 hours ago, Exoscientist said:

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.

To be clear, they're not using these engines on ANY first stage. A test hop of a second stage is still a test hop for a second stage, even if the first stage isn't used in the hop.

5 hours ago, Exoscientist said:

The speculation about using an aerospike or aero plug on the first stage . . . 

. . . which would not be possible due to thrust density problems . . . 

5 hours ago, Exoscientist said:

. . . is only relevant if they did use these upper stage engines on an operational first stage . . .

. . . which they aren't intending to do. 

5 hours ago, Exoscientist said:

even then it’s utility would be doubtful if the chamber pressure for these is in fact 100 bar . . . since you can get fairly high vacuum ISP with a hydrolox engine at only 100 bar as the Ariane 5 Vulcain demonstrates.

The Vulcain engine is a gas generator, not a bleed expander.

More importantly, the Vulcain is not powerful enough to lift even its own stage off the ground, because as I said above, it has a near-vacuum-sized nozzle, resulting in major thrust shortfalls at sea level. Without the solid boosters, the Ariane 5 would just sit there on the pad motionless until 70% of its first-stage propellant was expended. 

An aerospike design would be even worse.

[magnemoe]

6 hours ago, Exoscientist said:

 

  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

Here they have one or two turbopumps and lots of chambers and you control trust direction of rocket with valves to these chambers through a lot of pipes so not an place I would very high pressure. 

[sevenperforce]

32 minutes ago, magnemoe said:

Here they have one or two turbopumps and lots of chambers and you control trust direction of rocket with valves to these chambers through a lot of pipes so not an place I would very high pressure. 

Right, which further emphasizes why this would not be a good engine for a first stage. I don't believe the Soviets have ever gone above four thrust chambers for a single turbopump (not counting low-thrust verniers). If this engine design uses two turbopumps for 30 chambers and you'd need ~90 chambers to get off the ground with sufficient thrust for rapid reuse, then you're looking at an engine section with a turbopump assembly that is three times as high (in an already volume-bloated stage, due to the use of hydrolox instead of methalox) with a tangle of piping that would make even the most die-hard Cthlulu slashfic aficionados recoil in horror and revulsion.

There's a reason aerospikes never took off...particularly not for single-stick rockets.

Edited June 29, 2023 by sevenperforce

[magnemoe]

26 minutes ago, sevenperforce said:

Right, which further emphasizes why this would not be a good engine for a first stage. I don't believe the Soviets have ever gone above four thrust chambers for a single turbopump (not counting low-thrust verniers). If this engine design uses two thrust chambers for 30 chambers and you'd need ~90 chambers to get off the ground with sufficient thrust for rapid reuse, then you're looking at an engine section that is three times as high (in an already volume-bloated stage, due to the use of hydrolox instead of methalox) with a tangle of piping that would make even the most die-hard Cthlulu slashfic aficionados recoil in horror and revulsion.

There's a reason aerospikes never took off...particularly not for single-stick rockets.

Agree, also the Russians has the bells just below the pump, not long pipes and valves.
Not suited for first stage as its designed focus for second stage is to survive reentry from orbit. 

Yes its trendy now after Falcon 9 and Electron to have one vacuum version of main engine for second stage while first stage has an cluster of the same engines. 
This is not optimal, using an hydrogen engine for first stage is also not optimal as hydrogen is low trust. 

[sevenperforce]

5 minutes ago, magnemoe said:

its trendy now after Falcon 9 and Electron to have one vacuum version of main engine for second stage while first stage has an cluster of the same engines. 
This is not optimal, using an hydrogen engine for first stage is also not optimal as hydrogen is low trust.

Not only is hydrogen low trust, but it doesn't have great thrust either. 

Anyway you're right -- the trend of using clustered versions of the upper stage engine on the lower stage is recent. Before SpaceX, nobody ever did that (although you could kinda argue that Black Arrow and the Ariane 1/4 did it, but those are stretches). Since SpaceX, only RocketLab has done it. It's part of the plan for the US-based Relativity and ABL Space and for the Spain-based PLD Space, but none of them have gotten to orbit yet.

Using the same engine on the upper and lower stages and simply clustering can certainly add some simplicity, but it is by no means a secret sauce. 

[sevenperforce]

On 6/27/2023 at 12:06 AM, Exoscientist said:

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. 

On that note again -- what do you see as the advantage of placing a derelict payload into a useless orbit on a hacked-together vehicle (with an extremely weak first stage) that they never intend to commercialize, requiring an engine on the upper stage that you never intend to use more than once?

Is it bragging rights? "Woo, we've orbited a vehicle." Okay, great. Nobody is going to treat that as any particular evidence of the viability of your actual operational launch vehicle. You're going to need to do all the same work to demonstrate your operational vehicle which you would anyway. What do you gain?

How would you even propose they do this? It's unlikely that their second stage has a fully-loaded sea level TWR high enough to get off the ground on its own, let alone carrying an upper stage.

[Exoscientist]

 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

Edited July 6, 2023 by Exoscientist

[sevenperforce]

13 hours ago, Exoscientist said:

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.

How will it get off the ground? Its thrust capabilities are based on vacuum thrust off a second stage with a payload on top, not sea level thrust off the ground with both an upper stage and a payload.

13 hours ago, Exoscientist said:

With this booster stage having 30 engines, take a new smaller upper stage as having 3 to 4 engines. 

Which 3 to 4 engines?

Stoke currently has an open expander bleed engine with one turbopump and 15 chambers.

So they'd need to develop a new engine?

13 hours ago, Exoscientist said:

But what about the upper stage? . . . For it’s vacuum Isp I’ll take it as the RL10’s best 465.5 s.

So now they're buying an RL10? At $20-40 million?

To what end?

[Exoscientist]

On 7/6/2023 at 3:58 PM, sevenperforce said:

How will it get off the ground? Its thrust capabilities are based on vacuum thrust off a second stage with a payload on top, not sea level thrust off the ground with both an upper stage and a payload.

Which 3 to 4 engines?

Stoke currently has an open expander bleed engine with one turbopump and 15 chambers.

So they'd need to develop a new engine?

So now they're buying an RL10? At $20-40 million?

To what end?

 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

Edited July 8, 2023 by Exoscientist

[sevenperforce]

On 7/8/2023 at 9:20 AM, Exoscientist said:

 

On 7/6/2023 at 3:58 PM, sevenperforce said:

How will it get off the ground? Its thrust capabilities are based on vacuum thrust off a second stage with a payload on top, not sea level thrust off the ground with both an upper stage and a payload.

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.

400,000 feet is ~124 km, which would require about 1.45 km/s. Add a three-tonne upper stage and payload and I'm unsure that's enough umph to get high enough for the upper stage to achieve orbit.

On 7/8/2023 at 9:20 AM, Exoscientist said:

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.

I'm skeptical about such an architecture being able to get ANY satellites larger than cubesat size to LEO.

On 7/8/2023 at 9:20 AM, Exoscientist said:

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.

Well, the DC-X certainly was not going to be an SSTO.

Any SSTO which can reach orbit, deliver a payload, and return to the ground in one piece can deliver MORE payload with an upper stage.

On 7/8/2023 at 9:20 AM, Exoscientist said:

Quite notable as well is Bono wanted to use the efficiency of the aeroplug/aerospike for the launch, not disparage it.

Getting an SSTO or sustainer to orbit more or less requires an altitude-compensating nozzle system, yes. But you'll also note that none of Bono's designs proposed a using an aerospike on the first stage of a two-stage architecture. 

[Exoscientist]

 About the chamber pressure of the Stoke Space engines, I’m informed by someone who worked on the Everyday Astronaut web site that the 100 bar combustion chamber pressure that appears in the graphic explaining the expander bleed cycle was chosen arbitrarily to explain the cycle. It does not stem from something Stoke told them.

 Further evidence for this that this is that this chamber pressure of 100 bar also appears in an earlier YouTube video by Everyday Astronaut describing different rocket engine cycles when he explains the expander bleed cycle at about the 51 minute point in the video:

 

 Here’s, the graphic as it appears in this video:

 

 Note this video was made a year ago, so it was made before  Tim Dodd interviewed Andy Lapsa for the Stoke Space video.

 So that can’t be relied upon to conclude the Stoke Space engine is at 100 bar combustion chamber pressure. 

 But I did find further evidence that in fact the Stoke Space engines are at a low pressure of perhaps 40 bar, comparable to the RL10 upper stage engines. I saw this image on the Nasaspaceflight.com forum discussion thread on Stoke Space that was clipped from a Stoke Space video:

 At the upper left is given some pressure value of “40”, units likely being in terms of bar. I can’t read the subscript though. It could be turbine pressure or it could be combustion chamber pressure. Either way the chamber pressure is likely at or below 40 bar.

 Whether the chamber pressure is at 40 bar or 100 bar has important consequences. If at only 40 bar, then you really need large expansion ratio to get high vacuum ISP, needed for an upper stage engine. Andy Lapsa even alluded to this fact in the Everyday Astronaut video. But the key problem is at only 40 bar chamber pressure such high expansion ratios can not operate at sea level; the nozzles are dangerously over-expanded at sea level.

 But for landing you do want the engines to operate at sea level. So you would actually need  either an extensible bell nozzle like used on the highest Isp RL10 engine, or an aeroplug/aerospike to operate at sea level. There is no evidence Stoke Space will use extendable nozzles on the upper stage so it must be they will be using the aeroplug to get both high efficiency at vacuum and operability at sea level.

 This illuminates again why Stoke Space should mention that their idea of an aeroplug on the upper stage comes from Philip Bono, for Philip Bono also wanted to put an aero plug on that stage, which is needed to get good performance both at sea level and at vacuum. 

 So rather than disparaging the aerospike Stoke Space should acknowledge they need to use it.

 About the importance of knowing whether the engine is at 100 bar or 40 bar, note a hydrolox engine at 100 bar can get fairly good vacuum Isp of 434s while still being operable at sea level as proven by the Vulcain engine.

  Robert Clark

Edited July 11, 2023 by Exoscientist