Author Topic: ULA Innovation: Integrated Vehicle Fluids (IVF)  (Read 124291 times)

Offline Damon Hill

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Adding warm G02 and GH2 to tanks is not desireable is one is seeking low to zero boiloff.  Yes, GHe is eliminated per the documents.


The 'complaint' is that the system states all these grand goals, then compromises the IVF system architecture to meet a short term goal.


1.  warm G02 and GH2 is only added during engine operation where boil off is not a concern.

2.  Pray tell, how is the IVF system architecture "compromised"?
1. So this is not part of the overall future architectures...understood.  Why not cool the gas?
2. The pumps are mechanically driven.  So how are the pumps powered by Electric/Nuclear propulsion?

Learn to think flexibly.  IVF >is< flexible.

1.  There are short, medium and long duration missions; the IVF architecture being developed for Centaur will work very well for short to medium missions.  Cooling the gas is not appropriate for these missions.   Longer missions will need to take different approaches to conserving propellant so the specific implementation of IVF will be different.

2.  See #1.  I don't see any reason why for long duration missions, solar or nuclear power sources couldn't be used for sustained operations, and either a fuel cell or IC engine running off propellants couldn't handle relatively short peak power needs.

--Damon

Offline Jim

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Learn to think flexibly.  IVF >is< flexible.

1.  There are short, medium and long duration missions; the IVF architecture being developed for Centaur will work very well for short to medium missions.  Cooling the gas is not appropriate for these missions.   Longer missions will need to take different approaches to conserving propellant so the specific implementation of IVF will be different.

2.  See #1.  I don't see any reason why for long duration missions, solar or nuclear power sources couldn't be used for sustained operations, and either a fuel cell or IC engine running off propellants couldn't handle relatively short peak power needs.

--Damon

 IVF is only for cryogenic stages, specifically Centaur.  It is not generic term or idea applicable across other systems.


« Last Edit: 04/07/2015 06:00 pm by Jim »

Offline Damon Hill

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I was thinking within the range of chemical propulsion systems, which is what I thought this thread was originally about.

Offline russianhalo117

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I was thinking within the range of chemical propulsion systems, which is what I thought this thread was originally about.
As the creator of the topic thread, Jim is correct. There is additional published ULA white papers on this topic and more at: http://www.ulalaunch.com/Education_PublishedPapers.aspx

Offline baldusi

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Learn to think flexibly.  IVF >is< flexible.

1.  There are short, medium and long duration missions; the IVF architecture being developed for Centaur will work very well for short to medium missions.  Cooling the gas is not appropriate for these missions.   Longer missions will need to take different approaches to conserving propellant so the specific implementation of IVF will be different.

2.  See #1.  I don't see any reason why for long duration missions, solar or nuclear power sources couldn't be used for sustained operations, and either a fuel cell or IC engine running off propellants couldn't handle relatively short peak power needs.

--Damon

 IVF is only for cryogenic stages, specifically Centaur.  It is not generic term or idea applicable across other systems.
Could be used for CH4/LOX, thou. And they are going to move the first stage to that propellant with the BE-4. I'm not saying that they'll implement it on an upper stage. But Both BO and SpaceX could very well implement similar architectures in methane stages, if there are not patented technology.

Offline jongoff

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Basically, IVF takes a major liability of liquid hydrogen, and turns it into an asset.  The hydrogen and oxygen are going to boil off and be lost anyway, so use it for ullage, attitude control, pressurization and electric power.  Eliminate the hydrazine, high pressure helium and most of the batteries, which get heavy on extended duration flights.  IVF can increase payload by upwards of a ton, and increase flight duration to days instead of hours.

But its a giant piston engine with a bunch of steel in it. How is that lighter than anything else? How does using the IVF avoid having to use ullage thrusters? Still not understanding.

Why was hydrogen fuel cell chosen against?

I suggest you read articles posted earlier in this thread especially the patent link, it has most detailed information in it.

Yeah, I didn't see much steel in their piston engine prototype they were showing at NSS last year. It runs cool enough that I think most of the thing is aluminum.

~Jon

Offline Space Ghost 1962

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Basically, IVF takes a major liability of liquid hydrogen, and turns it into an asset.  The hydrogen and oxygen are going to boil off and be lost anyway, so use it for ullage, attitude control, pressurization and electric power.  Eliminate the hydrazine, high pressure helium and most of the batteries, which get heavy on extended duration flights.  IVF can increase payload by upwards of a ton, and increase flight duration to days instead of hours.

But its a giant piston engine with a bunch of steel in it. How is that lighter than anything else? How does using the IVF avoid having to use ullage thrusters? Still not understanding.

Why was hydrogen fuel cell chosen against?

I suggest you read articles posted earlier in this thread especially the patent link, it has most detailed information in it.

Yeah, I didn't see much steel in their piston engine prototype they were showing at NSS last year. It runs cool enough that I think most of the thing is aluminum.

~Jon

Modern ICE's for decades have been steel sleeves in cast aluminum blocks, few use cast iron.

Offline jongoff

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Basically, IVF takes a major liability of liquid hydrogen, and turns it into an asset.  The hydrogen and oxygen are going to boil off and be lost anyway, so use it for ullage, attitude control, pressurization and electric power.  Eliminate the hydrazine, high pressure helium and most of the batteries, which get heavy on extended duration flights.  IVF can increase payload by upwards of a ton, and increase flight duration to days instead of hours.

But its a giant piston engine with a bunch of steel in it. How is that lighter than anything else? How does using the IVF avoid having to use ullage thrusters? Still not understanding.

Why was hydrogen fuel cell chosen against?

I suggest you read articles posted earlier in this thread especially the patent link, it has most detailed information in it.

Yeah, I didn't see much steel in their piston engine prototype they were showing at NSS last year. It runs cool enough that I think most of the thing is aluminum.

~Jon

Modern ICE's for decades have been steel sleeves in cast aluminum blocks, few use cast iron.

This engine runs so rich (GOX/GH2 has an amazingly wide flammability range) that I think it might not even need steel sleeves, though I might be wrong.

~Jon

Offline Damon Hill

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Unlined aluminum blocks have been tried, but did not give satisfaction.  Anyone remember the original Chevrolet Vega engine?

The IVF development engine is rated for 26 HP, so it's not very big especially with the flathead construction.

I can almost hear the countdown status poll:

"Centaur IVF is go for engine start.   Vroom.  VROOM."

(pause)

"Centaur IVF, this isn't like driving the General Lee.  Stop racing the engine."

"Sorry.  Just kidding."

--Damon

Online TrevorMonty

1) what is the weight saving of IVF?

2) Do they plan to use 2 x ICE for redundancy?.

3) Is ICE used pre lift off to pressurise the Centuar tanks?.

Offline deltaV

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This engine runs so rich (GOX/GH2 has an amazingly wide flammability range) that I think it might not even need steel sleeves, though I might be wrong.

Do you know what mixture ratio it's designed for?

I believe Earth ICEs usually run pretty close to stoichiometric but they're kept cool by all the nitrogen in air. Even to match typical Earth ICE temperatures you'd need tons of unburned hydrogen so I'm skeptical of the no-steel-needed conjecture (but am not an engineer). It's certainly plausible that they might run it that fuel-rich so I'm not saying it's wrong, just I'd like to see more evidence before I'm convinced.

Offline Damon Hill

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Descriptions show two IVF platforms, presumably with two engines.  Electrical load and batteries are shared, as may be gases.  This helps insure both platforms continue have thrusters if there is an engine problem on one or if mission requirements at any given time needs only one engine running.

The IVF engine seems to be like a two-cycle engine with crankcase induction of hydrogen gas, direct oxygen injection to the individual cylinders, and exhaust through valves in the engine block rather than the header.  Any blowby (traces of O2, unburned H2 and water steam) into the crankcase is thus scavenged.  Lubrication is dry sump with centrifugal separation of gas and oil, like an aircraft engine.  Engine exhaust is used just like a rocket thruster for propellant settling; not quite as impulsive as an actual rocket thruster, but decent enough to be useful.

The engine will run hydrogen rich, but the documentation on the engine development suggests the oxygen ratio can be varied, and engine boost to double the horsepower was demonstrated, just like a supercharger.

I don't see any reason why IVF couldn't be operated as soon as propellant loading began, but I expect tank purge will be with ground He.  IC engines normally run heavily diluted with nitrogen, and He shouldn't be an impediment.  In general, it appears the engine will run from well before launch, during launch, and as long as necessary for orbit operations, payload deploy, and disposal maneuvers.  In short, a lot. 

Obviously, on the pad venting of hydrogen-rich engine exhaust would be required, and likely flared a bit remotely to prevent combustible buildups.

Somewhat detailed description of the engine development, explaining the particular choice of of engine design, is here:

http://tinyurl.com/ula-ivf2012

Not all operating parameters are detailed, as this is a development model, but it should be a general guide.

I have no idea if IC engines have ever been operated in space before, especially for extended periods.

--Damon
« Last Edit: 04/07/2015 11:22 pm by Damon Hill »

Offline Damon Hill

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1) what is the weight saving of IVF?


Weight savings through elimination of hydrazine, He, and most batteries have been suggested as 500 kg, but for longer duration missions with additional tankage and batteries thus no longer required, the savings should be even greater.  I've seen suggestions that payload increases could amount to a ton on some missions.  Better management of boiloff would contribute, too.

Presumably the final figures will vary a lot with specific missions and further development of systems.  It will be interesting to see how this translates to essentially free additional payload capability without any changes to the RL10 and the Centaur stage basic design.

Offline kevin-rf

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Sorry for the slightly off topic question, but with IVF would we still see super sync. GTO missions or would a Hohmann transfer with a apogee burn (now that stage life is not an issue) make more sense? My understanding (which may be faulty) is super sync is done because the stage has more delta V than needed for a strick Hohmann and really lacks the life (without kits) to do a partial circularization burn at apogee.

I may be completely off base.
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Online TrevorMonty

1) what is the weight saving of IVF?


Weight savings through elimination of hydrazine, He, and most batteries have been suggested as 500 kg, but for longer duration missions with additional tankage and batteries thus no longer required, the savings should be even greater.  I've seen suggestions that payload increases could amount to a ton on some missions.  Better management of boiloff would contribute, too.

Presumably the final figures will vary a lot with specific missions and further development of systems.  It will be interesting to see how this translates to essentially free additional payload capability without any changes to the RL10 and the Centaur stage basic design.
An extra 500-1000kg to GTO could save them millions on a mission by eliminating SRBs. The cost saves just keep adding up.
Most importantly this lower cost upper stage will be ready for NGLV booster.

With this new lower cost more capable upper stage as a stop gap ULA can afford to spend a few more years developing the ACES with new low cost engines.
« Last Edit: 04/08/2015 01:35 am by TrevorMonty »

Offline john smith 19

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Re-reading the report from 2012 in Damons' tinyurl entry reminds me just how small the thing is.

With no efforts at weight reduction the unit is 27 inches long and weights about 110lb while producing 26 Hp.

I think a lot of peoples intuition for the size of this thing when they hear "6 cylinder IC engine" is badly mislead by one simple fact.

It burns Oxygen not air. So whatever you think you need to burn air  is reduced to 1/5 that size.

Jongoff was right the block was Aluminum, with stainless steel exhaust manifold. This technology needed welded cap plates due machining certain blind channels in the block which would probably be viewed as unacceptable weak points in a production engine.

Does anyone know if complex Aluminum casting in small lots is fairly trouble free process or is this still only available in 10 000 unit lots with lots of fine tuning?
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You can 3D print one time use sand molds, ideal for small productions or prototyping. See video


Offline notsorandom

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Sorry for the slightly off topic question, but with IVF would we still see super sync. GTO missions or would a Hohmann transfer with a apogee burn (now that stage life is not an issue) make more sense? My understanding (which may be faulty) is super sync is done because the stage has more delta V than needed for a strick Hohmann and really lacks the life (without kits) to do a partial circularization burn at apogee.

I may be completely off base.
I'm speculating here without running the numbers. I don't think stage life is why super sync orbits are done, at least with ULA's rockets. One of the requirements of the EELVs (at least the Delta IV Heavy) was the ability to directly inject the payload to GEO. There are two benefits to dropping the payload off in a super sync orbit. The first is the benefit of staging. The Centaur masses a but under 3 mt dry and the 5 meter DCSS a bit over that. Unused propellants will make those figures higher. Though the stages have higher ISPs than the payload's propulsion sometime getting rid of that mass is the better way to go. The other thing is that in some situations a Hohmann transfer is not the most efficient. Plane changes take a lot of Delta V and there is a 28 degree one needed from the cape to GEO. The further from Earth that is preformed the less Delta V is needed. From the perspective of the total energy budget it can make sense to burn to a higher orbit then needed, do the plane change, then lowering the orbit.

Online LouScheffer

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Unlined aluminum blocks have been tried, but did not give satisfaction.  Anyone remember the original Chevrolet Vega engine?
An aluminum block might be fine here.  Even the Vega came with a 50,000 mile (~80,000 km) warranty.  At typical traffic speeds that's at least 1000 hours of operation.  This application needs only about 1/100 of that life.

Offline acsawdey

Unlined aluminum blocks have been tried, but did not give satisfaction.  Anyone remember the original Chevrolet Vega engine?
An aluminum block might be fine here.  Even the Vega came with a 50,000 mile (~80,000 km) warranty.  At typical traffic speeds that's at least 1000 hours of operation.  This application needs only about 1/100 of that life.

Porsche and others have apparently used aluminum blocks with a nickel silicon-carbide coating for quite some time:

http://en.wikipedia.org/wiki/Nikasil

A machined block is exactly the right solution for this application. The cost of machining is lost in the noise here and it'll be a lot easier to get the required level of quality assurance.

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