Author Topic: Rocket Engine Q&A  (Read 382943 times)

Offline fl1034

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Re: Rocket Engine Q&A
« Reply #760 on: 12/25/2019 12:43 pm »
Can someone tell me why LOX cooled kerolox ORSC engine has never been proposed by Aerojet Rocket dyne even after end of the cold war and availability of the RD180 which clearly shows the potential of an high ISP and thrust US indigenous engine? Switching to LOX cooling would completely eliminate the coking problem of RP-1.

Offline Unrulycow

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Re: Rocket Engine Q&A
« Reply #761 on: 12/25/2019 01:20 pm »
Can someone tell me why LOX cooled kerolox ORSC engine has never been proposed by Aerojet Rocket dyne even after end of the cold war and availability of the RD180 which clearly shows the potential of an high ISP and thrust US indigenous engine? Switching to LOX cooling would completely eliminate the coking problem of RP-1.
It has. The AR1 is an ORSC engine proposed for use in Vulcan.  They choose BE4 instead and AR1 may potentially be used in Firefly Beta in the future.
https://en.m.wikipedia.org/wiki/AR1

Offline fl1034

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Re: Rocket Engine Q&A
« Reply #762 on: 12/28/2019 11:29 am »
Can someone tell me why LOX cooled kerolox ORSC engine has never been proposed by Aerojet Rocket dyne even after end of the cold war and availability of the RD180 which clearly shows the potential of an high ISP and thrust US indigenous engine? Switching to LOX cooling would completely eliminate the coking problem of RP-1.
It has. The AR1 is an ORSC engine proposed for use in Vulcan.  They choose BE4 instead and AR1 may potentially be used in Firefly Beta in the future.
https://en.m.wikipedia.org/wiki/AR1

Well, why not LOX cooling? Splitting the usual coaxial shaft pump of the kerolox engine into 2, use LOX and LOX only for regen cooling, adjust the fuel side preburner temperature, then you get completely independent mixture ratio control, no coking to worry about, much easier pumps to design, and the ability to use from methane all the way to kerosene or even diesel for fuel. Just saying.

Offline Redclaws

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Re: Rocket Engine Q&A
« Reply #763 on: 12/28/2019 12:37 pm »
The obvious concern - which may not be relevant, but still - is well, it’s LOX.  It’s comparatively *very* nasty, even cold, and is going to introduce new material requirements in the cooling channels.  I would be interested to hear livingjw on this topic.

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Re: Rocket Engine Q&A
« Reply #764 on: 12/28/2019 01:02 pm »
I recall reading an old ('70s) study on LOX cooling on NTRS, which found that, surprisingly, this was not as dangerous as you might imagine, even if the cooling channels sprung a leak.
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Offline john smith 19

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Re: Rocket Engine Q&A
« Reply #765 on: 12/28/2019 01:12 pm »
Well, why not LOX cooling? Splitting the usual coaxial shaft pump of the kerolox engine into 2, use LOX and LOX only for regen cooling, adjust the fuel side preburner temperature, then you get completely independent mixture ratio control, no coking to worry about, much easier pumps to design, and the ability to use from methane all the way to kerosene or even diesel for fuel. Just saying.
That is a very good question.

It also means you don't have to fiddle with the cooling circuit if you change fuels, but not oxidizer (and given LOX is about the best available in terms of performance and cost why would you?)

In fact it has been used in at least one engine (the rocketdyne plug nozzle test bed for the USAFRL in the mid 70's used a dual expander cycle with LOX cooling and LH2 cooling of the modular combustion chambers to drive the separate pumps, eliminating interpropellant seals and the criticality 1 failure modes. 

Likewise Rotary Rocket tested LOX cooling in the early 90s. NASA ran tests on a 40 000 lb pressure fed test engine (late 80's, early 90'x) which included deliberate  leaks into the chamber. Nothing bad actually happened.

Here's the thing.  The bulk of rocket engineering was done in the 50's and 60's. It was done at breakneck speed (by modern standards) because of the Cold War.  If something looked too difficult to deliver in the short term it was discarded.

And TBH many of those decisions have never been reviewed or retested.

So generations of folklore and myth have accumulated (passed on by lecturers who've either never worked with it or never been asked to question what they have been told by their lecturers) that "Oh noes. LOX will burn anything"
It's quite interesting that AFAIK the propulsion team at Rotary had not gone through the conventional aerospace engineering education process and so had not received the stories, myths and old wives tales that have built up. 

The other classic meme is that HTP in unstable, despite it being specified for "Super performance" booster engines on early jet aircraft and supplying on orbit station keeping for a comms sat for 6 years in the early 60's.

IRL LOX cooling is likely to cool the area around the hole and reduce the risk of ignition. A big enough leak in the chamber cooling system is a serious issue regardless of what the coolant is.
« Last Edit: 12/28/2019 01:13 pm by john smith 19 »
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Offline john smith 19

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Re: Rocket Engine Q&A
« Reply #766 on: 12/28/2019 01:18 pm »
The obvious concern - which may not be relevant, but still - is well, it’s LOX.  It’s comparatively *very* nasty, even cold,
Welcome to the site.

Actually it's nasty mostly because it is cold.  If it leaks it warms up, vaporizes and disperses. Contrast that with NTO, which is seriously toxic as a liquid and makes a quite good WMD if it vaporizes. It has also been know to explosively decompose.
Quote from: Redclaws
and is going to introduce new material requirements in the cooling channels.  I would be interested to hear livingjw on this topic.
I'd be interested in hearing more from HMX on the subject, given his company actually built LOX cooled combustion chambers.

Or you could just look up the NASA papers on www.sti.nasa.gov for the subject.
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Offline john smith 19

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Re: Rocket Engine Q&A
« Reply #767 on: 12/28/2019 01:26 pm »
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage? 
IIRC both design teams had it as their US engine.

Quote from: darkenfast
Didn't the people who pushed the SSME for the Upper Stage know this? 
Well you'd think so but actually Rocketdyne was asked a similar question in the early 90's (IIRC) and said it wouldn't be a problem.  IIRC it's mostly the augmented spark ignitors that didn't have enough flow.
Quote from: darkenfast
How were they going to deal with this?
Well they'd been told it wasn't a problem.

You might think with both designs relying on the SSME doing a start in space that NASA would have either a) Request the teams demonstrate this or b)Scheduled some stand time on a altitude test stand to verify it.

But they did neither.

It's been noted before that NASA is much more willing that the DoD to go ahead with ideas with much lower levels of proof that they will even work.

I'm not sure how many $Bn were spent before this issue was finally "discovered." :(
« Last Edit: 12/28/2019 04:42 pm by john smith 19 »
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline envy887

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Re: Rocket Engine Q&A
« Reply #768 on: 12/30/2019 03:39 pm »
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage? 
IIRC both design teams had it as their US engine.

Quote from: darkenfast
Didn't the people who pushed the SSME for the Upper Stage know this? 
Well you'd think so but actually Rocketdyne was asked a similar question in the early 90's (IIRC) and said it wouldn't be a problem.  IIRC it's mostly the augmented spark ignitors that didn't have enough flow.
Quote from: darkenfast
How were they going to deal with this?
Well they'd been told it wasn't a problem.

You might think with both designs relying on the SSME doing a start in space that NASA would have either a) Request the teams demonstrate this or b)Scheduled some stand time on a altitude test stand to verify it.

But they did neither.

It's been noted before that NASA is much more willing that the DoD to go ahead with ideas with much lower levels of proof that they will even work.

I'm not sure how many $Bn were spent before this issue was finally "discovered." :(

Wasn't the main issue lack of pressure head in freefall after staging? I'm curious why they didn't opt for large solid ullage motors, or hot staging, as those would provide the acceleration needed for a pressure head.

Online DaveS

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Re: Rocket Engine Q&A
« Reply #769 on: 12/30/2019 05:50 pm »
All good facts, but then: wasn't the SSME the first choice for Aries I Upper Stage? 
IIRC both design teams had it as their US engine.

Quote from: darkenfast
Didn't the people who pushed the SSME for the Upper Stage know this? 
Well you'd think so but actually Rocketdyne was asked a similar question in the early 90's (IIRC) and said it wouldn't be a problem.  IIRC it's mostly the augmented spark ignitors that didn't have enough flow.
Quote from: darkenfast
How were they going to deal with this?
Well they'd been told it wasn't a problem.

You might think with both designs relying on the SSME doing a start in space that NASA would have either a) Request the teams demonstrate this or b)Scheduled some stand time on a altitude test stand to verify it.

But they did neither.

It's been noted before that NASA is much more willing that the DoD to go ahead with ideas with much lower levels of proof that they will even work.

I'm not sure how many $Bn were spent before this issue was finally "discovered." :(

Wasn't the main issue lack of pressure head in freefall after staging? I'm curious why they didn't opt for large solid ullage motors, or hot staging, as those would provide the acceleration needed for a pressure head.
Airstarting the SSME after staging wasn't the problem, it was re-starting it for the TLI burn. NASA wanted a common engine for both the Ares I and Ares V upper stages so when they ran into the problem of getting a good re-start of the SSME on the Ares V upper stage, they decided it wasn't worth expending the resources into making it re-startable in flight and went with the J-2X as the upper stage engine on both LVs. This had the cascade effect of making the Ares I upper stage unable to insert the CEV into orbit which forced them to not only upsizing the upper stage but the booster stage as well (was derivative of the STS SRB but became a brand new five segment version with no flight history).
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Offline john smith 19

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Re: Rocket Engine Q&A
« Reply #770 on: 12/30/2019 08:57 pm »
Airstarting the SSME after staging wasn't the problem, it was re-starting it for the TLI burn. NASA wanted a common engine for both the Ares I and Ares V upper stages so when they ran into the problem of getting a good re-start of the SSME on the Ares V upper stage, they decided it wasn't worth expending the resources into making it re-startable in flight and went with the J-2X as the upper stage engine on both LVs. This had the cascade effect of making the Ares I upper stage unable to insert the CEV into orbit which forced them to not only upsizing the upper stage but the booster stage as well (was derivative of the STS SRB but became a brand new five segment version with no flight history).
Wow.

That is a truly impressive lack of anything like foresight.
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Offline nicp

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Re: Rocket Engine Q&A
« Reply #771 on: 01/03/2020 09:26 pm »
What is the big deal about throttling liquid fueled rocket engines? This was definitely considered a big deal for the LEM descent engine, and even that (which clearly worked well) could get erosion at certain throttle settings.
As I understand it _that_ engine displaced some of the propellants with inert helium.

But surely, say, Merlin 1D - which has got to have a quick and deep throttle to land - how is that done?
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Offline john smith 19

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Re: Rocket Engine Q&A
« Reply #772 on: 01/06/2020 08:41 pm »
But surely, say, Merlin 1D - which has got to have a quick and deep throttle to land - how is that done?
It isn't.

F9 takes off on 9 engines then lands on 1.

That alone deals with most of the "deep" throttling problem.

A Merlin at 50% thrust (a fairly shallow throttle providing your nozzle expansion ratio is not extreme) has 1/18 of an F9's take off thrust.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline walkermo

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Re: Rocket Engine Q&A
« Reply #773 on: 01/13/2020 09:13 pm »
What are the challenges around designing an electric pump engine similar to the Rutherford or Stealth Space/Astra’s engine? How much complexity & cost does it add over a pressure fed system?

Asking as a member of a student team that has successfully built and tested a N2O/Kerosene pressure fed engine.

Offline TrevorMonty

Re: Rocket Engine Q&A
« Reply #774 on: 01/13/2020 10:55 pm »
What are the challenges around designing an electric pump engine similar to the Rutherford or Stealth Space/Astra’s engine? How much complexity & cost does it add over a pressure fed system?

Asking as a member of a student team that has successfully built and tested a N2O/Kerosene pressure fed engine.

Design suitable battery, power electronics, electric motor and pumps. Maybe able to source motor off shelf but battery pack would be custom build. Keeping everything cool at these high power levels would also add to design challenges.

Would be good project for student rocket but would take few years and multi discipline team. Great to have on team members' CV even if not successful as all these disciplines are in high demand across lots of industries.
« Last Edit: 01/13/2020 10:56 pm by TrevorMonty »

Offline Hog

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Re: Rocket Engine Q&A
« Reply #775 on: 01/14/2020 05:14 pm »
But surely, say, Merlin 1D - which has got to have a quick and deep throttle to land - how is that done?
It isn't.

F9 takes off on 9 engines then lands on 1.

That alone deals with most of the "deep" throttling problem.

A Merlin at 50% thrust (a fairly shallow throttle providing your nozzle expansion ratio is not extreme) has 1/18 of an F9's take off thrust.
I remember the paper dealing with the RS-25 17%RPL 24%RPL 40%RPL deep throttle testing.  It's amazing how complicated things cab get.
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Offline robert_d

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Re: Rocket Engine Q&A
« Reply #776 on: 06/04/2020 08:42 pm »
 I had a question that is sort of related to the development of the Super Heavy booster by SpaceX. I thought, especially before the engine count reduction to thirty-one, that stuffing them all into nine meters was impossible so that they would need a larger ‘skirt’ or shroud to shield the sides of the outer ring. It appears to be a good idea to keep that ring as small as possible. I also know that certain existing rocket engines such as Atlas 5’s RD-180 have one set of turbo machinery feeding two engine bells. So is the reverse possible? Could multiple engine sets feed into one bell to save space? What I conceived was a compound shaped bell with three lobes that had an exit area close to the exit area of three standard engines. This would need to assume that the turbo-machinery could coexist nearer to each other than any limitation of the outer bell diameter. Is this at least theoretically possible? I have attached a picture of the conceptual shape I had in mind. The outer triangle depicts a hemispherical divide where the normal engine bells would just touch. The inner triangle defines the position where the total exit area of the lobed shape would approximate the actual area of three standard engine bells.
Anyone care to comment?

Offline robert_d

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Re: Rocket Engine Q&A
« Reply #777 on: 06/04/2020 08:57 pm »
... Could multiple engine sets feed into one bell to save space? What I conceived was a compound shaped bell with three lobes that had an exit area close to the exit area of three standard engines. ...
Clarification: This would all occur AFTER the actual combustion chambers and hopefully only deal with the geometry of the final exhaust and cooling thereof.

edit: "and cooling thereof" isn't clear but I was trying to acknowledge that the extra-ordinary shape might cause complications in any regenerative cooling system.

edit 2: note on outer circle in the diagram - it isn't anything except the it was the only way I could find to hold the proper geometric center of the lobed shape in my inexpensive CAD program.
« Last Edit: 06/07/2020 12:06 am by robert_d »

Offline Crispy

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Re: Rocket Engine Q&A
« Reply #778 on: 06/04/2020 09:45 pm »
If one of your three combustion chambers suffers a failure, then you have severely unbalanced pressure inside the bell. It would likely deform or suffer from flow separation on the faulty side. You'd have to shut all three chambers down to avoid further damage. This reduces engine-out capability

You have less surface area to absorb the same amount of energy via bell cooling. This reduces efficiency (heat is "wasted" by ejecting a hotter exhaust)

If you really want packing efficiency, you might consider hexagonal bells, but no pressure vessel likes being far from a circle without getting heavy (a downside that especially applies to the lobed shape)

Offline robert_d

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Re: Rocket Engine Q&A
« Reply #779 on: 06/07/2020 12:24 am »
If one of your three combustion chambers suffers a failure, then you have severely unbalanced pressure inside the bell. It would likely deform or suffer from flow separation on the faulty side. You'd have to shut all three chambers down to avoid further damage. This reduces engine-out capability


Thanks for your thoughtful response.
I can understand the first objection but wonder if the fact that since a standard engine can throttle down to 66% or more this might not be a deal breaker. All changes would be reductions in pressure and the flow would not be affected until some point beyond the combustion chamber. So can I hold to a faint hope that it wouldn't be as bad as you think? 

 

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