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#51 2020-08-17 16:17:39

GW Johnson
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From: McGregor, Texas USA
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

From post 49:  "For the second, vacuum part of the flight you should have used the Isp of the vacuum optimized engine, ca. 380s."

WRONG! 

I was NOT analyzing any sort of altitude compensating nozzle!  I was analyzing the fixed bell fixed geometry engine as currently is flying.  It is the sea level Raptor with only 40:1 expansion,  operated in vacuum,  and its performance in vacuum is 356 s Isp,  per Spacex. 

The vacuum-optimized engine has 200:1 expansion,  not 40:1 expansion.  That 200:1 is what gets you 380 s Isp in vacuum.  It cannot be operated below about 60,000 feet because of backpressure-induced separation.  Such CANNOT be used for surface-launch SSTO!

I've already looked at free-expansion nozzles for altitude compensation.  From the surface to some design altitude in the atmosphere,  they equal or slightly exceed the thrust and Isp performance of an equivalent-thrust fixed bell,  even one designed at 20,000-to-30,000 feet.  From there into vacuum,  their performance steadily falls way below fixed bell performance,  due to streamline divergence expanding too much into vacuum.  And that vacuum fixed-bell performance is the for that same sea level or not much higher expansion,  not what you might do in a "vacuum-optimized" design.

There is one and ONLY ONE way known to get fixed bell performance near sea level,  and also some sort of "vacuum-optimized" fixed bell performance in vacuum.  And that is some sort of variable geometry that adds a bell extension,  once a certain critical altitude is exceeded,  so that there is no backpressure separation risk.  NONE SUCH are flying!

There have been experiments,  but no such technology is actually ready to fly yet.  The cooling problem is nearly insuperable,  and plugging the gas leaks at the joint is very nearly as insuperable.  Both problems are fatally catastrophic.  Which is PRECISELY WHY none such are currently flying!

My purpose here,  as was stated,  was to look at TSTO vs SSTO with technologies "already in-hand".  I did NOT look at what might be done with a new technology requiring considerable development before you can even begin to trust it.  Only in that case would I use the "vacuum-optimized" Isp performance nearer 380 s.  But I would NOT know how to build it.

One final note:  there is NO SUCH THING as a "optimized vacuum expansion" for a fixed bell,  despite the widely-misused terminology.  There is only what fits the base of your stage!  No two such designs are alike.  There is NO accepted standard for the specific value expansion ratio that is called "vacuum-optimized".  For Spacex's Raptor it is 200:1.  In my old Pratt & Whitney Handbook,  it is 100:1. 

Everybody uses different values.  Which is WHY quoting "vacuum Isp" is BS,  there is only delivered chamber c* and what you do with it.

A truly "vacuum-optimized" bell would have an infinite expansion ratio,  at an infinite exit diameter.  That is because you must expand to Mach number infinity in order to have an expanded static pressure of exactly zero to match your ambient vacuum.  That's the definition of "perfectly expanded".  But,  so also will the expanded (absolute) temperature be zero!  There are no materials in the universe that stay gases at a static temperature of zero!  So it is quite fundamentally a BS concept.

GW

Last edited by GW Johnson (2020-08-17 16:30:12)


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#52 2020-08-17 18:24:48

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

kbd512 wrote:

Bob,
Increasing the first stage / burn Isp to 360s and the second stage (TSTO) or second burn (SSTO) Isp to 380s won't produce a significantly different result than what GW already provided.  The payload performance of the TSTO is at least 4X better for the same 2nd lowest performance propellant combination.  Increasing first burn Isp by 4.7% and the second burn Isp by 5.5% is NOT going to make a vehicle that's 4 times heavier a cost-competitive solution.  The inert mass fraction of the LCH4 SSTO is 84,240kg.  The inert mass fraction of the LH2 SSTO is 17,550kg...

This is actually the crux of the matter. Even a 10% increase in Isp can result in a 100% increase in payload for a SSTO. That is the nature of exponential change. Most people interested in space have heard the phrase, “the tyranny of the rocket equation.”
The other side of the coin is a small change in Isp can result in a radical change in payload. I like to call this “the beneficence of the rocket equation.”

  Bob Clark

Last edited by RGClark (2020-08-17 18:46:41)


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

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#53 2020-08-17 18:30:10

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

GW Johnson wrote:

From post 49:  "For the second, vacuum part of the flight you should have used the Isp of the vacuum optimized engine, ca. 380s."
WRONG! 
I was NOT analyzing any sort of altitude compensating nozzle!  I was analyzing the fixed bell fixed geometry engine as currently is flying.  It is the sea level Raptor with only 40:1 expansion,  operated in vacuum,  and its performance in vacuum is 356 s Isp,  per Spacex. 
...
GW

OK. But for a SSTO to be most efficient altitude compensation is needed. I’m aware that no alt. comp. nozzles are currently flying. My point is a radical increase in payload would be possible for a SSTO if they were.

Edit: to be more precise there are no extensible nozzles firing from sea level. There are extensible nozzles on the upper stage engine RL-10B2.

rl10b2.jpg

But it would be possible to do to make the extensions operate from sea level up to vacuum. You could have multiple extensions on a sea level version. See for example how small the nozzle is on the version used on the low altitude DC-X test vehicle:

RL10A-5.jpg

  Bob Clark

Last edited by RGClark (2020-08-17 18:45:31)


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#54 2020-08-17 21:43:22

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Bob:

To do an extensible nozzle for the LOX-LCH4 designs that I analyzed,  you would need precisely the sea level Raptor at 40:1 expansion,  with a bell extension that would extend it to 200:1 expansion.  I have no clue how to build that,  or if the extension geometry is even compatible with the powerhead. 

But if such a thing were possible (and it might be),  then you follow the performance curve for the sea level 40:1 engine with increasing altitude up to that altitude where you shut off thrust long enough to deploy the extension.  Then you relight,  and you follow the performance curve for the same powerhead running with the 200:1 expansion. 

One needs to "optimize" the altitude at which the extension is deployed.  That is the only way I see to fly into vacuum at 380 sec Isp instead of 356 s Isp.  It's just that nobody knows how to actually do it.  Not even Spacex,  or they would have already done it.

BTW,  I saw a posting tonight that confirms Spacex has finally reached and slightly exceeded the design 300 bar chamber pressure of the Raptor powerhead.  Bravo,  and those 330/356 (and 380) s Isp numbers are now finally "real"!  Those numbers are lower at throttled-back lower chamber pressures.  I figured them from ballistics and posted them over at "exrocketman" some time ago,  in my reverse-engineering article about the Raptor.

The same thing would be true of the LOX-LH2 RS-25,  whose data I used to evaluate those designs.  The exponential variation of mass ratio and Isp favored that propellant combination over LOX-LCH4 for SSTO operation.  I think the figures I found illustrate that outcome quite clearly.  If some variable geometry bell extension could be had for the RS-25,  then yes,  its vacuum Isp could be higher.  And that would indeed make the SSTO performance more attractive. 

But I really think you need gas core nuclear thermal to make the SSTO truly more attractive than any possible chemical TSTO.  And it really would be.  Probably with the 1300 s Isp of the light-bulb engine cycle,  if you can accept the higher inert fraction imposed by the lower engine T/W (better than NERVA,  but still far less than chemical).  It would buy you a non-radioactive exhaust plume with a nuclear engine.  But you still have a very radioactive core on board after shutdown. 

And,  I think you can achieve a very attractive SSTO with that kind of engine.  But once again,  no light bulb engines have ever actually been built,  much less flown.  Not even NERVA ever flew,  although it was ready to fly when Nixon's order killing manned flight outside LEO effectively killed it.  That's just where we are,  technologically.

The only real difference here,  between what you have been saying,  and what Kbd512 and I have been saying,  is a technology readiness thing.  You are oriented toward what could be,  if the supporting technology were ready.  We are oriented toward what actually is,  with the technology that is actually in hand.  Both viewpoints are valid,  and both are needed.  Such guides what technologies need to be matured.

The trick with bell extensions is how much extra expansion,  and at what altitude to deploy it.  Thrust and Isp curves vs altitude are the way to determine that.  But those curves need to be traceable and "real",  to be believable. 

If I can ever get my laptop back from the shop repaired,  I will get my nozzle spreadsheet to you.  You can use it to get those curves with altitude,  because that is exactly what I set it up to do.  It handles any propellant combination!  All you need to know is the effective value of the product gas specific heat ratio.  Which is never very far from 1.20.  The ballistics require an estimate of delivered c* as a function of chamber pressure,  and for the reality check, you need an estimate of what fraction of total gas generated gets dumped overboard running the turbopumps.

GW

Last edited by GW Johnson (2020-08-17 21:56:43)


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#55 2020-08-18 16:22:08

SpaceNut
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

The RL10 series is very versatile flight-proven hydrogen/oxygen expander cycle engine.
https://en.wikipedia.org/wiki/RL10

https://ntrs.nasa.gov/archive/nasa/casi … 010379.pdf

Even though I believe testing was done with many other fuels beyond those that are production runs use.

https://www.benjaminmunro.com/liquid-ox … evelopment

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#56 2020-08-21 09:06:13

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

GW Johnson wrote:

...
If I can ever get my laptop back from the shop repaired,  I will get my nozzle spreadsheet to you.  You can use it to get those curves with altitude,  because that is exactly what I set it up to do.  It handles any propellant combination!  All you need to know is the effective value of the product gas specific heat ratio.  Which is never very far from 1.20.  The ballistics require an estimate of delivered c* as a function of chamber pressure,  and for the reality check, you need an estimate of what fraction of total gas generated gets dumped overboard running the turbopumps.
GW

I got the spread sheet. Thanks. It’s very useful for doing estimates. You might want to make it available on Google Docs so anyone interested in doing the calculations of engine performance with various chamber pressures and expansion ratios can do them.

I have a question about the example “Conventional 3”. It discusses a fixed nozzle, I assume, kerolox engine at a 1800 psi, 120 bar, chamber pressure and 35.5 nozzle area ratio. This results in a surprising high 341 s vacuum Isp. This would mean a sea level fireable engine reaching the high vacuum Isp of the Merlin Vacuum upper stage engine.

However, the spreadsheet notes the flow separation would be close to happening at sea level. Likely though a slightly higher chamber pressure would make this less likely to happen. This is interesting because I did a calculation using GW’s criterion in post #28 for when flow separation could occur applied to the high performance Russian kerosene engines the RD-180 and RD-191, and was surprised to see their chamber pressures were so high at ca. 260 bar, that you could put larger nozzles on them to get 360s vacuum Isp yet would still be fireable from sea level, i.e., no flow separation.

If this is correct then these engines all you would have to do is put longer fixed nozzles on them and you would have SSTO capable engines, no altitude compensation required. This is surprising if true. Because the higher vacuum Isp would be important to get higher payload for the standard TSTO as well. So why didn’t the Russians give the engines the larger fixed nozzles to begin with? 

This example “Conventional 3” does have a somewhat higher chamber pressure at 120 bar compared to the Merlin’s 100 bar, though. It would be interesting to do the calculation to see how much payload the “Falcon 9” could get to orbit with this engine instead of the Merlin’s. Note also with this fixed nozzle it could be SSTO with significant payload, no altitude compensation required.

I put “Falcon 9” in quotes because this is a much larger nozzle size than the Merlin so 9 wouldn’t fit underneath the same diameter booster. So you would have to design engines with higher thrust so a fewer number of them could lift the same size vehicle.

Also, note that looking at example “Conventional 1” with the smaller nozzle area ratio 13.8 to 1. This is close to the Merlin 1D sea level engine’s nozzle area ratio of 16 to 1. But its sea level Isp is much higher at 303 s compared to 282 s. And it’s vacuum Isp is also better than the Merlin 1D at 324 s compared to 311 s.

This illustrates the usefulness of alt comp. It could improve the sea level Isp to 303 s by using the sea level area of 13.8 to 1 while giving the 341 s vacuum Isp by having the adaptive nozzle expand to 35.5 to 1 in vacuum.

   Bob Clark

Last edited by RGClark (2020-08-21 09:09:56)


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#57 2020-08-21 09:35:36

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

I have a question also in the spreadsheet about the axial aerospike #2 example. It has a sea level Isp of 308 s and a vacuum Isp of 363 s. It gives the aerospike length as 45”, a little more than 1 meter. This for an engine of about 3,000 lbs thrust. How long would the spike be to match the Merlin thrust of about 200,000 lbs thrust?

  Bob Clark


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#58 2020-08-21 16:28:09

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Bob:

The right answer from me about liquid engine designs is "I dunno".  Somebody with real liquid experience would know more. 

But,  there are a host of issues limiting what you can do.  We have been discussing backpressure-induced flow separation,  which leads to (1) massive loss of thrust and Isp,  and (2) probably fatal heating issues if allowed to go on for any significant amount of time.  We have also discussed how the vacuum expansion ratio is limited by how much bell diameter will fit behind the stage. 

In addition to those,  there is how much bell length will fit in the interstage between stages,  which I think is the genesis of the RL-10B configuration for which you showed the illustration. 

And there is loss-of-thrust at sea level with a bell that has not quite separated.  That is because the exit plane pressure minus backpressure is a large-magnitude negative number,  which must be multiplied by a large exit area.  That negative number "adds" to the exit velocity x massflow rate momentum thrust term.   For launch applications,  a thrust that low may not be at all tolerable,  when the vehicle is so heavy,  and the aerodynamic drag so significant.

And,  there is the ignition thrust level issue known since Von Braun's experiences at Peenemunde.  You cannot successfully light an engine at max thrust propellant flow rates.  It blows up.  You MUST light at reduced flow rates,  which is reduced thrust and chamber pressure conditions.  How much reduced?  I dunno,  a real liquid guy might.  But during that interval,  your nozzle is separated because the backpressure/chamber pressure ratio is too high.  This won't correct until you successfully throttle up.  How long can you survive the adverse heating during this transient? Aye,  there's the rub,  and I don't know the answer.  But it ain't very long.  I do know that.

OK,  for the free expansion spike nozzles,  the spike length is proportional to the throat diameter.  Look at the throat area data in F = CF At Pc for a guide.  Each area has a diameter.  The spike length to throat diameter would be constant.

GW

Last edited by GW Johnson (2020-08-21 16:29:46)


GW Johnson
McGregor,  Texas

"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#59 2020-08-22 23:50:02

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

GW, I’m interested in any and all methods of altitude compensation, from the aerospike, to methods of physically changing the nozzle size, to any other method.

One interesting possibility is to not vary the nozzle size at all but to vary the throat size, keeping in mind it’s the nozzle area to throat area ratio that needs to vary.

So I was interested to read this comment of yours:

Advanced Aerospike Rocket Engine Design
These tactical nozzle designs included both straight conical and full curved bell designs.  We never tried anything else as a production item.  The flapper-lollipop throat insert as a variable-geometry nozzle for ASALM is the exception.  It worked,  but the real-world design restrictions were twofold:  (1) the wake zone behind the lollipop had to close before the exit plane if full nozzle efficiency was to be obtained,  and (2) you could not put a large pressure drop across a structure like that because there were (and are) no materials capable of withstanding such abuse.
http://newmars.com/forums/viewtopic.php … 80#p146380

Could you expand on the problems you observed with that?

   Bob Clark


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#60 2020-08-23 12:27:46

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

That was an option for a variable-geometry nozzle for the ASALM-PTV ramjet test vehicle.  That was a prototype for a supersonic cruise missile that we never built.  It did fly as a target drone called SLAT for USN.  But both SLAT and the test vehicle flew with fixed nozzle geometry only.  No lollipops at all.

Throat area was smaller with the lollipop sideways to the flow,  mounted in the throat plane.  Throat area was around 2-3 times larger with the lollipop streamlined.  This was at ramjet pressures,  which never ever exceeded about 200 psia,  and were usually below about 50 psia.  It could not survive rocket pressures (1000-2000 psia) and by orders of magnitude. 

It was held streamline in boost,  ahead of the booster nozzle insert that was ejected at transition from rocket to ramjet.  The rocket booster propellant was cast within the ramjet chamber.  The inlet duct was sealed off during boost by a frangible glass port cover. Boost was only about 3 seconds long,  the real thermal protection problem was ramjet sustain.  The hot gas washed over this thing with sonic speeds at a chamber temperature very nearly 4000 F.  We used steel substrate covered in silica phenolic,  and got it to survive 900 second burns,  and still actuate,  even though it really was ugly-looking by that time.

For rocket-level pressures,  you would be far better off inserting a rounded-end cylindrical pintle from the side across the throat.  We did that very successfully at 2000-2500 F gas temperatures as the throttle valve on the solid propellant gas generator of a gas generator-fed ramjet.  For that application,  we were able to use bare metal components made of TZM,  because the stream flowing by it was reducing,  and TZM has a meltpoint close to 4700 F.  However,  TZM rapidly corrodes away if oxygen is present above about 1300 F. We easily got throat area ratios equalling or exceeding 7:1. 

For a liquid propellant rocket,  you won't be able to use bare TZM,  the flame temperatures are just too high at 5000-6000 F.  You may still require TZM instead of more ordinary high-alloy steel,  but you will have to coat the outside of it with an ablative like silica phenolic.  You will be able to insert the pintle further as time goes by,  but you will NOT be able to retract it,  because the silica phenolic char will swell its effective diameter too large to return back into the hole from whence it came.

Fortunately for you,  you need to insert the pintle further to get the higher expansion ratio with a fixed bell.  It will work for ascent,  but you are just screwed for landing!  You will NOT be able to retract the pintle to get a sea level expansion ratio for landing.  It will scrape the charred silica phenolic right off the steel substrate. This is NOT a reusable thing.  You have to replace the pintle insulation for every flight,  and (like I already said) it is only good for ascent.

You try to stick bare metal into the sonic point of a 6000 F stream,  and the liquid cooling requirements are going to be utterly unbelievable.  Far,  far,  far worse than the throat of a conventional bell. A bare metal pintle you can extend and retract will require manurium,  unbelievium,  or maybe some unobtainium.  There are simply no real-world materials that will serve your purpose.  We got away with it on the gas generator-fed ramjet because of the lower flame temperatures of a very fuel-rich gas generator system.

GW

Last edited by GW Johnson (2020-08-23 12:35:01)


GW Johnson
McGregor,  Texas

"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#61 2020-08-23 15:33:41

kbd512
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

GW,

Hafnium Carbide melts at 7,156F.  Tantalum Hafnium Carbide melts at 7,619F, though most variations on this ceramic metal melt closer to 7,200F.  Both were intended to be used in sharp leading edges of hypersonic vehicles.  Sufficient to withstand 6,000F, maybe?

Tantalum Hafnium Carbide is about $9,500/kg, so I hope whomever is paying for this miracle rocket nozzle material has very deep pockets.  Even if you only needed 100kg per engine, you're adding a million dollars to the cost of each engine.  There had better be a way to recycle this stuff, too.  It's not unobtanium in the strictest sense of the word, but pretty close.  Maybe Uncle Sam can afford small quantities of this stuff for hypersonic weapons or the hot sections of jet engines, but there are no non-governmental aerospace applications that I'm aware of.

I'm aware of ceramic adhesives that can withstand temperatures up to 4000F or so, but no more than that, which begs the question of how these components would be attached or bonded to the rest of the engine to dump their thermal load into the propellant.  To the best of my knowledge, these UHTC aerospace components are small "thermal soak" parts wherein the sharp leading edge heats up and dumps the heat generated into a refractory metal skin structure / fuel tank heat sink combination.  NASA / USAF / DLR have experimented with hypersonic vehicles using this technology.  I think our military jet engines now use some kind of UHTC in their exhaust nozzles and burners as well, but that's it.  I don't think any rocket engine nozzles use this technology, so there may be some problem with using it that way.

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#62 2020-08-25 12:32:18

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

That's the problem with all the ultra-high temperature ceramics:  how do you hold onto them? 

They are all high-density parts.  Which means high thermal conductivity.  They will inherently equilibriate at very nearly isothermal soak (constant temperature) throughout. 

If the part is near 6000 F at one end,  it will be near 6000 F at the other end.  So,  what do you use to connect to it?

We got the miracle material.  What is still missing is the supporting technology required to actually use it.

GW


GW Johnson
McGregor,  Texas

"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#63 2020-08-29 09:26:22

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

GW Johnson wrote:

...
For rocket-level pressures,  you would be far better off inserting a rounded-end cylindrical pintle from the side across the throat.  We did that very successfully at 2000-2500 F gas temperatures as the throttle valve on the solid propellant gas generator of a gas generator-fed ramjet.  For that application,  we were able to use bare metal components made of TZM,  because the stream flowing by it was reducing,  and TZM has a meltpoint close to 4700 F.  However,  TZM rapidly corrodes away if oxygen is present above about 1300 F. We easily got throat area ratios equalling or exceeding 7:1. 

...GW

So were you able to get a functioning altitude compensating ramjet that way?

   Bob Clark


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#64 2020-08-29 14:09:12

kbd512
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

GW,

I guess you'd have to cool the rest of the structure using the cryogenic propellants.  Since that works for hypersonic vehicles designed to fly for half an hour or longer, as that was part of their design criteria (the sharp leading edge flight experiments, SHEFEX I / II / III, DLR conducted that I've posted about in other threads), why wouldn't that same cooling method work in a rocket engine that's only operated for a few minutes at a time?

I must be missing something about this.  The leading edges of those flight test vehicles (they've flown several times now using rocket propulsion to get them up to speed, in addition to many hours of wind tunnel testing) were subjected to extreme aerodynamic heating for extended periods of time and the metallic skin (high grade nickel steels, mostly) were bolted to the ceramic parts, why is this not possible to do with a rocket engine?  Why wouldn't this work better when the high temperature components aren't subjected to such temperatures for one to two hours at a time?  If the materials weren't withstanding the temperatures or oxidation environment acceptably well, why did DLR continue on with their experiments?

DLR and Airbus are working towards a hypersonic high altitude air-breathing aircraft for commercial passenger transport service across the oceans to other continents.  If you combine sharp leading edges with Skylon's engine technology, then you have a practical hypersonic transport aircraft that can deliver people between continents in one or two hours instead of the better part of a day.  The vehicle takes off and lands like a conventional aircraft on conventional runways, so the notoriously dangerous VTOL flight regime is not required.  SpaceX's approach may well prove more practical and it will get you there faster, but it won't compete on fuel costs.  Fuel and maintenance are what drive ticket prices, as you already know.  Perhaps SpaceX's approach will cost less in the end if the engine maintenance is minimal (I can't see how that will be the case for an engine that operates at Raptor's chamber pressures, though), at the expense of a greatly increased fuel burn.

Skylon is ~325t and carries 30 passengers to orbit, but if we only have to cruise at Mach 5.5 at 85,000 feet, then it could feasibly carry a lot more passengers than for an orbital flight.  Given the tonnage of LOX carried, I see no reason why you couldn't eliminate the rocket engines and LOX tanks to drastically simplify the vehicle and its propulsion system and carry as many passengers as a standard wide body.  A 20 hour flight would be over with in 3 hours or so.  The orbital version is in the same MTOW class as the Boeing 777-300ER, but it's about 40 feet longer and the wings are much shorter.  I'm much more interested in a practical intercontinental transport that's 5 times faster than it's competitors than a payload-limited SSTO, even if Skylon can theoretically use 1/5th as much propellant as a conventional rocket by breathing air.  The complexity of everything involved for SSTO to work is far higher than an advanced high speed transport.

Edit:
To be perfectly clear, I think SpaceX will have a lock on the orbital transport business, but Skylon will have a lock on the intercontinental passenger transport business and pose far less risk to the flying public and general public than gigantic VTOL rockets the size of office buildings.

Last edited by kbd512 (2020-08-29 14:18:16)

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#65 2020-08-30 15:39:33

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Bob & Kbd512:

The side-inserted pintle was a very-well-tested throttle valve used with the fuel-rich solid propellant gas generator of a gas generator-fed ramjet.  It did not soak out too hot,  even after a minute-and-a-half long burn.  We used graphite rings around the pintle shaft to act as a heat sink. 

It was a "one shot design" that we refired (with refurbishment) many,  many times.  "Refurbishment" means replacing the silicone O-ring gas seal at the cool end of the pintle,  the silica-phenolic gas passage insulator shells,  and any worn metal parts. We tested this thing in subscale lab motor hardware,  then graduated to full-scale AMRAAM-class hardware.  It is the fuel control in the AGM-163 "Coyote" gas generator-fed ramjet target drone.

While the gas generator flame temperatures were only 2000-2500 F,  bear in mind that most fuel propellant effluents are a two-phase mix of gases and soot plus some oxide particles.  30-50% solids in the effluent was pretty common.  Those solids were also at full chamber temperature,  by the way. Very harsh environment,  indeed.

This valve was used with a very non-linear control system as an arbitrary command throttle control.  The early linear controls were failures,  it had to be very nonlinear,  with an adaptive gain,  in order to control the very nonlinear ballistics of the solid propellant generator. 

When I say "failure",  I mean the motor blew up.  We NEVER blew up one with the nonlinear control! Not in 15 years of testing!

It not only did altitude compensation,  it did arbitrary command fuel flow between the flowrate limits for the generator-valve system.  For most of the tests,  we had 7+ flowrate turndown ratio available between end-burning generator pressures of 1500-2000 psia hot to 50-70 psia cold.

This was military stuff.  Hot meant 145 F,  cold meant -65 F. With an irreversible variable-area grain design,  you could get even more turndown ratio.  We usually just went ahead and designed with the older mil std "hot temperature" of 165 F. Desert ammo piles have been observed even hotter than that.

For the constant area end-burner,  you could command max or min flow anytime you wanted,  during the burn.  Or anything in between.  Arbitrary fuel flow profiles vs time were what this thing provided!  You needed solid propellants with a burn rate law exponent about 0.7 to make this work right.  No less than 0.55,  no more than 0.85.

I also did a no-moving-parts throttle based on a French design,  which did altitude and speed compensation,  but not hot/cold soak compensation. That turned out to be the safest and fastest way to do experimental lab tests of multiple fuel propellant formulations for the gas generator-fed ramjet. 

That one used a large unchoked throat for the gas generator,  so that its internal pressure was the forward-end ramjet combustor pressure.  That way the ramjet pressure drove the fuel propellant burn rate,  reflecting how much air was being scooped up,  in turn dependent upon altitude and speed.

If the burn rate law exponent was the (usually unconditionally-unstable) unity,  compensation was "perfect".  This generator acted like a big strand bomb at ramjet chamber pressure.  The French didn't have fuel propellants with exponents near unity;  we did.  That lack caused poorer performance than they wanted,  which is why they never fielded it. We were going to,  but got closed.  I tested it many times in full-scale AMRAAM-class hardware.

For varying rocket nozzle throat area to vary expansion ratio with a fixed bell,  you want the choked side-inserted pintle.  It'll take some materials and heat transfer smarts to modify the old design for clean gas at up to 6000 F.  For burns only a couple of minutes long,  you might get away without active cooling. Longer,  and I bet you have to remove heat from the graphite heat sink "real time". Maybe not all of it,  but a lot of it.  Even so,  that should be "do-able".
 
GW

Last edited by GW Johnson (2020-08-30 15:40:20)


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#66 2020-09-01 01:18:54

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Thanks for that. For the reusable rocket application, the “no-moving-parts throttle“ looks to be the way to go to avoid the extreme heating issues in the throat for a rocket engine.

SpaceX also wants high throttle ability for landing, so perhaps you could provide them advice there:

https://twitter.com/search?q=Throttle%2 … ery&f=live

  Bob Clark


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#67 2020-09-01 08:24:26

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

I don't think you want the big unchoked throat as a rocket engine throttle.  It is inherently a low chamber pressure system,  which is "death" for thrust force out in the vacuum off space.  There is no bell acceleration capability in principle,  because nowhere does the flow Mach number ever reach 1. It worked as a ramjet fuel supply control because the air massflow dominated,  and its velocity dominated even more,  in all of the combustor processes. Getting adequate mixing and distribution of fuel mixture was more difficult with it. NONE of that obtains in a rocket engine!

That is why I suggested the side-inserted pintle that worked so well in the arbitrary commanded-fuel system.  It is choked,  the throat plane has the pintle centered in it,  and that is the Mach 1 point.  As you change the throat area from max to min,  several things happen:  (1) the exit/throat area ratio factors up by the same factor that throat area factors down,  (2) in a liquid system,  at constant chamber pressure and c*,  the total flow rate factors down by the same ratio that the area factors down,  and (3) thrust reduces,  but not by the factor that ratios throat area. 

The only downside I can think of is that nozzle efficiency is no longer just the streamline divergence factor.  Based on the lollipop we tested in the ASALM ramjet nozzle,  you factor together about .983 for a 15 degree (average) bell and about .96 or .97 for about a 3-4% total pressure loss through the shock field behind the pintle,  for a nozzle efficiency of about 0.94 or 0.95.  That's just a guess,  but we found numbers like that in ASALM.  We didn't test that in the fuel throttle application,  because fuel thrust force wasn't the issue,  controlled massflow delivery was.

GW

Last edited by GW Johnson (2020-09-01 08:27:25)


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#68 2020-09-12 19:03:45

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Bob:

To follow up,  I looked at a couple of side-inserted pintle nozzle designs.  One was sized to Spacex's sea level Raptor conditions,  the other to Spacex's vacuum Raptor conditions.  They fell in between the sea level and vacuum Raptor designs in performance,  but will operate from ses level to vacuum,  like the sea level Raptor (the vacuum design cannot do that).  The shortfalls in performance for the variable geometry designs trace mainly to the additional shock loss factor on nozzle efficiency.  There is a wake behind the pintle,  an expansion off-centerline downstream of it,  and a shockdown back to parallel flow,  all of which must occur before you expand in the final bell.   

My best shot was a bit lower in SL performance at about 320 s Isp,  and a bit short in vacuum performance at about 372 s Isp.  I got the same throat diameters and sea level flowrates as Spacex did with their designs.  My designs run constant Pc with decreasing At by reducing flow rates as the pintle gets inserted further.  SL Raptor is 330-ish s Isp at SL,  and 360-ish s Isp in vacuum.  The vacuum Raptor is about 381 s Isp in vacuum,  but cannot be operated below 20+ kft altitude at 4400 psia,  and cannot be operated below 50+ kft altitude at 20% (880 psia).

Can't say more.  I may file a patent on this thing.

GW

PS 9-13-20:  further followup.

Spacex supposedly has a long-bell Raptor capable of ground test in the open air,  at something barely short of flow separation.  It's not a "vacuum-optimized" engine (in point of fact there is no such thing),  but it does supposedly have an expansion ratio in the vicinity of 120,  according to things I have seen online.

I looked at that.  About 92.5 is as high as I could go in expansion without violating my own criterion for separation,  and even that requires one to operate within 1 or 2 % of max Pc = 4400 psia.  What that means is that the online stuff about 120 expansion ratio is just BS.  You cannot go that high,  and expect it to work right.  My separation criterion may be conservative,  but it is not inaccurate.  That bell design showed low sea level thrust (not surprising due to the larger backpressure term) of nearer 400 than 440 klb,  at about 320 Isp.  Vacuum thrust was good,  with Isp about 372 s.  Bad for takeoff at heavy weight,  good for high final acceleration toward orbit.

It's Murphy's Law / TANSTAAFL thing.  If you improve vacuum performance in terms of Isp,  you must pay for it elsewhere with reduced thrust and /or Isp,  usually with lousier sea level performance.  Make sea level performance better,  and you WILL lose vacuum performance. 

Or else you will be heavy,  complex,  and/or unreliable.  Sealing the extendible bell can be done,  but getting it done with reliable sealing is a lot of trouble.  If the sealing fails,  the bell is destroyed (and that usually means loss of vehicle,  one way or another).

You cannot win.   There is NO solution that gets near-optimal performance at all altitudes from sea level to vacuum.  You WILL pay a price to get any improvement anywhere along the trajectory in one engine.  THAT is why two stages with different engines (or at least different bells) is the most popular solution.  It really is the most practical thing to do. 

The side-inserted pintle nozzle loses a little sea level Isp to gain some (but not all) the vacuum Isp,  but it does so at greatly-reduced high altitude thrust levels (sea level thrust is not reduced).  From a trajectory dynamics standpoint,  that may,  or may not,  be an acceptable thing to incur.  If it is acceptable,  then the weight of the throttle gear might be worth it.  But nowhere along the trajectory would the Isp look like vacuum levels. 

The only other thing I didn't look at was maintaining flow rate as throat area reduces,  and accepting the higher operating chamber pressures.  Pc is already 4400 psia in the Raptor.  Another 4 to 10 times higher than that is just ridiculous.

And all of that is why I think getting vacuum-level performance all along a trajectory from sea level to space is just unattainable,  even in principle!  Down in the air,  the backpressure term on thrust will always reduce your Isp and thrust,  or else it will limit your free expansion,  or semi-free expansion,  to lower exit Mach numbers.  That's just physics.  It cannot be sidestepped. --  GW

PS Update 9-18-20: 

Further investigation reveals that the proper way to use this variable throat area device is NOT to get perfect expansion ratio during the ascent,  but as a simple two-position geometry change at much higher altitudes, up near the TSTO stage point. That is because when you reduce throat area at constant chamber pressure,  you also reduce flow rate and thrust.  You CANNOT TOLERATE a thrust reduction until you reach the TSTO stage point on a gravity turn trajectory,  where the path is very nearly horizontal,  the weight is far lower,  and the Wsin path angle term is very small.  Up there,  drag is essentially zero. Low thrust actually works.   

So you run it with wide open throat and high thrust at sea level expansion ratio,  until you are exoatmospheric and nearly horizontal,  somewhere around 150 kft.  Then you shut down briefly,  reduce the throat to min,  and fire back up at low thrust and flow rate,  but at vacuum-type expansion ratio and Isp.  From there,  you thrust to orbit.  It's a single engine with two expansion ratios,  in a single bell that is about the size of a typical sea level bell.    You pay for this capability (TANSTAAFL):  it will be heavier,  due to the throttle gear,  and due to needing either two sets of turbopumps,  or a very complex turbopump,  to handle two wildly-different flow rates to the same high delivery pressure.   The higher the vacuum Isp you attempt to get,  the higher the throat area turndown ratio must be,  and lower that high-expansion thrust level will be.   Which pushes the transition to higher altitudes where the path is even flatter.  The practical limit is about 10:1 area turndown.

-- GW

Last edited by GW Johnson (2020-09-17 08:35:01)


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#69 2020-09-18 00:42:51

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Thanks for that. Perhaps we can reduce the loss of efficiency due to shock loss by carefully shaping the pintle that inserts into the throat? Perhaps to something similar to the cone shape seen inside the inlet of a ramjet or scramjet?

About the SpaceX Raptor version at 120 to 1 expansion ratio, I mentioned before in post #56 I did a calculation on the famous RD-180 class of Russian kerolox engines, that seemed to show with their high ca. 260 bar chamber pressures, by changing to a larger nozzle, it could with that single nozzle be operable at sea level while being able to get ca. 360 s vacuum Isp, i.e., no flow separation with the larger nozzle at sea level due to the high chamber pressure.

Perhaps you could try that type of calculation on the Raptor? How high would the chamber pressure have to be to be able to use a single nozzle and get ca. 330 s sea level Isp and ca. 380 s vacuum Isp. Actually because of the higher chamber pressure it would probably get even higher sea level Isp.

I like your idea of just two variable area settings. This would be analogous to having an extendable bell attachment like the RL10-B2 engine.

  Bob Clark


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#70 2020-09-19 09:03:07

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

Bob:

Did you get my email with the performance plots?  They were for sea level Raptor,  vacuum Raptor,  and an intermediate-expnsion design that I investigated.  Should answer most of the questions you posed.

I don't think streamlining the side-inserted pintle would help with the shock-loss effect.  The sonic line is at the min area throat.  Flow past that around the pintle is supersonic,  an inherently off-axis toward the centerline as the pintle wake closes.   Streamlining just fills the wake,  the off-axis flow toward the centerline is inherently still there.  Being supersonic,  when the flow straightens out to be axial,  there must be oblique shock waves.  That's just compressible flow physics. 

I'm amazed the shock field does not cause more loss than it does.  I used a total pressure ratio number similar to what we saw in actual tests with the ASALM lollipop design.  It had the same wake closure problem behind the lollipop as exists behind the pintle.  It is just that the side-inserted pintle gets you a lot more area turndown (on the order of 10:1) in a geometry that can be structurally survivable,  than the lollipop ever could (around 3 or 4).

GW


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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#71 2020-09-25 01:30:52

RGClark
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

...
PS 9-13-20:  further followup.

Spacex supposedly has a long-bell Raptor capable of ground test in the open air,  at something barely short of flow separation.  It's not a "vacuum-optimized" engine (in point of fact there is no such thing),  but it does supposedly have an expansion ratio in the vicinity of 120,  according to things I have seen online.

I looked at that.  About 92.5 is as high as I could go in expansion without violating my own criterion for separation,  and even that requires one to operate within 1 or 2 % of max Pc = 4400 psia.  What that means is that the online stuff about 120 expansion ratio is just BS.  You cannot go that high,  and expect it to work right.  My separation criterion may be conservative,  but it is not inaccurate.  That bell design showed low sea level thrust (not surprising due to the larger backpressure term) of nearer 400 than 440 klb,  at about 320 Isp.  Vacuum thrust was good,  with Isp about 372 s.  Bad for takeoff at heavy weight,  good for high final acceleration toward orbit.

It's Murphy's Law / TANSTAAFL thing.  If you improve vacuum performance in terms of Isp,  you must pay for it elsewhere with reduced thrust and /or Isp,  usually with lousier sea level performance.  Make sea level performance better,  and you WILL lose vacuum performance. 
...
-- GW

Here’s a video clip of the Raptor Vacuum being tested on the ground:

https://twitter.com/erdayastronaut/stat … 26240?s=21

At a 107 to 1 expansion ratio, it has a shorter nozzle than the expected operational version. Someone in that twitter discussion suggested it had a support ring placed around the nozzle to deal with the high side loads.

  Bob Clark

Last edited by RGClark (2020-09-25 02:20:37)


Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):

      “Anything worth doing is worth doing for a billion dollars.”

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#72 2020-09-25 08:17:01

Oldfart1939
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

There appears to be an external support ring visible.

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#73 2020-09-25 08:51:03

GW Johnson
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Re: Starhopper+Starship for heavy. Triple-cored Starship for super heavy.

I don't know whether that's a structural support ring,  or an added cooling manifold,  or both. 

You can very easily see the over-expanded flow condition.  The plume boundary tucks inward sharply,  then more-or-less straightens-out axial at the first big shock diamond. 

The nozzle is on the verge of backpressure-induced separation,  but it is not separated.  If it were,  you would see a plume not as big as the bell exit originating from somewhere up inside the bell.  You would also likely see a hot spot on the bell just downstream of the separation point. 

These things are a whole lot easier to see in liquid engines with an all-gaseous stream.  In the solids,  you could not see clearly,  because of the glowing condensed phases obscuring everything,  especially with metallized propellants.  With some of the reduced-smoke (non-metallized) formulations,  we could see the shock diamonds fairly clearly.  Usually,  I could see the shocks and flow processes in the gas generator-fed ramjets,  despite the solids, except for the magnesium-fueled versions.  Those were opaque white.

When the flow in a nozzle separates,  there is a more-or-less conical shock wave system set up at the separation point,  similar to (but not identical to) the leading edge of a typical shock diamond.  This thing at least triples the boundary layer thickness,  making it very unstable.  Once it forms at all,  this migrates further upstream to an equilibrium position in a perceptual instant.  There is very strong amplification of the heating rates just downstream of the shock,  which forms the hot spot on the bell.  This is not always symmetrical in shape.

The purported 107 expansion ratio is not far at all from the 92.5 I found using my rough-and-ready separation criterion.

GW


GW Johnson
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"There is nothing as expensive as a dead crew,  especially one dead from a bad management decision"

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