You are not logged in.
Playing with NASA rocket simulator, I "discovered" that NTO-MMH rockets have not the optimal O/F ratio: it is kept near 1-1.2 to have an acceptable chamber temperature near 2500 °K, but the optimal ratio, which gives the best Isp, is about 2.4. Using this O/F ratio the vacuum exhaust velocity rises over 3600 m/s (with an expansion ratio of about 500) but the drawback is that even the chamber temperature rises over 3200°K.
Anyway, in Orbitec vortex-cooled rockets, the oxidizer is injected tangentially to the chamber, creating a vortex that keeps the combustion in the middle without touching the wall.
http://www.celestialmechanics.co.uk/vortex.html
Combining these two items, we can make high Isp storable propellant rockets, ideal for a Mars manned mission, because NTO-MMH tanks can be easily sent to Mars orbit as return propellant with electric propulsion, without boil-off trouble.
Last edited by Quaoar (2018-10-29 16:15:33)
Offline
Like button can go here
Quaoar,
I presume the stability of this type of combustion process is well-characterized, since this looks like an application of combustion instability to prevent the effects of combustion instability from destroying the engine. Would that be correct, or is this technology still in development?
It looks very interesting, though. I'm definitely going to read more about it.
Offline
Like button can go here
Quaoar,
I presume the stability of this type of combustion process is well-characterized, since this looks like an application of combustion instability to prevent the effects of combustion instability from destroying the engine. Would that be correct, or is this technology still in development?
It looks very interesting, though. I'm definitely going to read more about it.
Orbital Technology Corporation (Orbitec) still produces commercial vortex cooled rocket
https://www.sncorp.com/press-releases/s … ex-rocket/
This is the takeoff of a vortex cooled rocket
https://newatlas.com/orbitec-vortex-liq … 7/#gallery
Last edited by Quaoar (2018-10-30 01:35:43)
Offline
Like button can go here
With most liquid propellant combinations, peak delivered Isp in tests does NOT occur at the stoichiometric oxidizer/fuel ratio by mass “r”. It peaks a little, but not a lot, on the fuel-rich side.
The following is what my Dec 1969-vintage Pratt & Whitney Aeronautical Vest-Pocket Handbook has for typical storables.
The sea level cases are for Pc = 1000 psia (feeding the nozzle), and the vacuum cases are for Pc = 100 psia (yes, only 100 psia!!!!). Sea level is perfect expansion to 14.7 psia, with 100% nozzle kinetic energy efficiency, and a shifting equilibrium composition all the way to the exit.
Vacuum is arbitrary perfect expansion to a 40:1 exit/throat area ratio, shifting equilibrium to the exit, and 100% nozzle kinetic energy efficiency.
Most realistic fixed-geometry nozzle designs test at about 98% nozzle kinetic energy efficiency (self-compensating and variable-geometry designs test lower !!!!). Isp is “seconds”. The only one of these combinations to have an r near 1-ish instead of near 2-3 (or more) is NTO-plain hydrazine. Everything else is pretty much r = 2+, with NTO-UDMH approaching 3.
You get more realistic estimates yet, with real kinetic efficiency efficiencies, and with shifting composition chamber to the throat, followed by frozen composition throat to exit. Small effects, but real. Sort-of 1-2% differences from what the handbook gives.
Case…………………....r…………SL Isp…………r……….vac Isp
NTO-MMH………..…2.17…..288…………..2.26…..338
NTO-Aerozine50….2.0……..287………….2.0……..339
NTO-UDMH………….2.6……287…………..2.7….….336
NTO-hydrazine……..1.33….292………….1.36……342
IRFNA--RP-1………….5.0……263…………..5.1……..309
IRFNA-UDMH……….3.1…….272………….3.2………320
Pure H2O2-UDMH..2.27….288………….2.32…….338
Pure H2O2-B2H6….1.84…..333………….1.86…….391 (see below!!!!)
Unfortunately, this old handbook does not list high-test H202 with RP-1. I’ve been told a vacuum Isp is about 305 sec with this. Be aware that raising Pc raises Isp, and this is a rather significant effect. It is due to both higher nozzle thrust coefficient, and higher thermochemical chamber temperature. Both are significant effects.
Note that Aerozine-50 is 50%-50% hydrazine-UDMH. IRFNA is about 83% nitric acid with 2% water, and 15% NTO. All of the various hydrazines are similar in toxicity and handling characteristics to the same anhydrous ammonia that farmers handle all the time, except that the vapor pressures are lower than ammonia. Toxicity lists as “medium”, just like ammonia.
NTO is an extreme toxic agent (lists as “high”): small exposures indeed are untreatably fatal. IRFNA, while toxic, is far less toxic that straight NTO, just a strong corrosive, with potential to convert ordinary oxygenated and non-oxygenated hydrogen-carbon compounds to explosives, similar to plain nitric acid. It lists as “medium”.
Most liquid propellant field weapons (like Scud and Lance) were fueled by the IRFNA-kerosene combination. That was the safest and most practical field handling for real troops, not expressly-trained technical specialists.
RP-1 is a refined kerosene, highly-filtered, with its own appropriate trace additive package, but otherwise similar to JP-5, JP-8, K-1, and Jet-A. These list as “low” toxicity. Chemical formulas for those who want to investigate stoichiometric r:
Item………………………….formula…..toxicity
NTO…………………………..N2O4……..high
Nitric acid………………….HNO3……..medium
Hydrazine………………….N2H4……..medium
MMH…………………………CN2H6…..medium
UDMH……………………….C2N2H8…medium
Diborane……………………B2H6 (very highly toxic, extremely expensive, not used for much anymore)
Kerosene…………………..C1H2 empirical, with effective molecular weight exceeding 150
The same ancient 1969 reference has things like boil and freeze points, specific gravities, vapor pressures at 77 and 160 F, viscosities, costs-per-pound in 1969 dollars, and a crude toxicity ranking (categories of “nontoxic”, “low”, “medium”, and “high”). I always found it a pretty good reference. Later years, not as good, lacking these details.
Don’t jump on diborane as a forgotten-to-death possibility! It proved very difficult to achieve anything close to theoretical performance out of any boron-containing compound, mostly due to the 4500+ R (2500+ K) melt temperature of the condensed-phase boron oxide product. Between that, the high toxicity, and the cost, just about all the boron compounds were not practical in any sense of the word. The main exception being solid BKNO3 used in some solid rocket igniter assemblies.
I did find a way to burn boron efficiently in a ramjet engine (NOT a rocket!!!!), but it’s not what or how you might think. I found other things just as good, or better, and that included one fuel with no metal at all!!!! None of these were liquid fuels.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
GW-
In the chemical industry these days, Hydrazine is viewed as Highly Toxic due to long term after effects on the liver; easily absorbed through the skin by contact. This applies equally to Aerozine-50, MMH, and ADMH. As a chemical professional myself, these are not problematic to handle if normal Personal Protective Equipment (PPE) is properly employed. NTO falls into the same category w/r PPE. My experience is there's nothing too dangerous to handle by skilled professionals, properly equipped and trained.
The major issue with hydrazine toxicity is carcinogenicity. That spooks a lot of people.
Offline
Like button can go here
OF:
My ancient reference is too old to reflect the modern views on toxicity. Typical at NASA they use full suits with self-contained breathing gear to handle the NTO-hydrazine systems. Actually getting some NTO in your lungs was usually an immediate death sentence. It was a worse threat than the hydrazines, which were bad enough.
Soldiers in the field without such gear were able to handle IRFNA-kerosene without much problem. That's a lower-performing system, but it works, and with hypergolic ignition. Such missiles were common in the 1950's, but pretty much got replaced by the "wooden-round" solid propellant systems by the 1970's. Far easier to handle. The still-lower performance was not objectionable for those applications.
I've been told that H2O2-kerosene is not really hypergolic, except maybe at H2O2 purity near 100%. Some claim stability for high test H2O2, but its reputation was poor in prior decades. Lots of stuff blew up, including one USN submarine. Even today, no one is claiming storage times longer than a few weeks for high-test H2O2. Most only claim days.
IRFNA isn't "just" nitric acid, there is some NTO in it. But it's mostly just a strong corrosive that must also be isolated from cellulose to prevent nitration into something explosive. It's probably easier to handle in the field than LOX. Certainly far less complex to handle.
GW
Last edited by GW Johnson (2018-10-31 08:10:52)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
OF:
My ancient reference is too old to reflect the modern views on toxicity. Typical at NASA they use full suits with self-contained breathing gear to handle the NTO-hydrazine systems. Actually getting some NTO in your lungs was usually an immediate death sentence. It was a worse threat than the hydrazines, which were bad enough. [...]
Soldiers in the field without such gear were able to handle IRFNA-kerosene without much problem. That's a lower-performing system, but it works, and with hypergolic ignition. Such missiles were common in the 1950's, but pretty much got replaced by the "wooden-round" solid propellant systems by the 1970's. Far easier to handle. The still-lower performance was not objectionable for those applications.
I've been told that H2O2-kerosene is not really hypergolic, except maybe at H2O2 purity near 100%. Some claim stability for high test H2O2, but its reputation was poor in prior decades. Lots of stuff blew up, including one USN submarine. Even today, no one is claiming storage times longer than a few weeks for high-test H2O2. Most only claim days.
IRFNA isn't "just" nitric acid, there is some NTO in it. But it's mostly just a strong corrosive that must also be isolated from cellulose to prevent nitration into something explosive. It's probably easier to handle in the field than LOX. Certainly far less complex to handle.
GW
Hi, GW
I used NASA software to simulate a real rocket like the AMBR (NTO-MMH, O/F 1.2, chamber pressure 18.96 bar, expansion ratio 1:400)
with these parameters it gives a vacuum exhaust velocity of 3340.7 m/s versus 3283 m/s of the real rocket. So the real rocket is about 98.3 % of the ideal simulated rocket.
Changing the O/F ratio to 2.4 and enhancing the nozzle expansion ratio to 1000, the simulator gives an exhaust velocity of 3722.2 m/s: taking the 98.3% of this value, a real rocket with these parameters should have an exhaust velocity of 3658.9 m/s: do you think we can thrust the simulator?
Last edited by Quaoar (2018-10-31 10:08:07)
Offline
Like button can go here
A factor which has discouraged more widespread use of NTO, Hydrazine, and IRFNA, has more to do with transportation costs and EPA regulations than the efficiency of them as fuels/oxidizers. They are all pretty easy substances to manufacture, but very difficult to transport to point of use. Probably the least problematic is IRFNA. In the 1960s Hydrazine had a very low cost to use as fuel, but these days it's in the hundreds of dollars per gallon to purchase and deliver.
Offline
Like button can go here
As GW pointed out earlier, Hydrogen Peroxide, H2O2, is a poor choice for a storable oxidizer; decomposition over even very short intervals makes it both inefficient and dangerous. The German Me 163 rocket fighter used it in combination with Hydrazine, but not what I would call successfully. Many operational accidents occurred, both in flight and in ground fuelling operations.
Offline
Like button can go here
To answer Quaoar in post 7, yes, I think your Isp simulator software is doing a decent job, as far as it goes. I think you might find a nozzle with an area ratio = 1000 is not something easily built, and likely rather heavy.
OF: H2O2 was used as a decomposition propellant in the launch tracks that flung the V-1 Buzz Bomb to launch speed. When the USN added their copies of V-1 as "Loon" to some submarines, they used solid rocket motors instead to achieve near-zero-length-launch from a short ramp behind the conning tower. Much lower risk. This was ca.-1950 stuff.
SS X-1 was a small submarine one-of-a-kind built in 1957 and retired in 1973 by the US Navy. It is on display at Annapolis. It had peroxide-diesel propulsion, and suffered an explosion that very nearly destroyed the boat and its crew. It was our shot at a peroxide diesel design inspired by the 1944-vintage Type XX U-boat design that was never produced.
The 1945-vintage Type XXI was a super-streamlined diesel battery design that was otherwise similar to the earlier Type VII and Type IX U-boats, just far more effective, and the inspiration for our GUPPY streamlining program. The safety problems with peroxide didn't stop the Nazis, it was inability to produce so much peroxide so fast that did stop them.
All: in my ancient handbook, LOX-UDMH and LOX-hydrazine look every bit as good as LOX-LCH4 in terms of Isp. If you can stand the boiloff of LOX, that's a semi-storable combination. If boiloff is not tolerable, you're back to NTO and the hydrazines.
GW
Last edited by GW Johnson (2018-10-31 14:33:17)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
Or nitric acid ...
Offline
Like button can go here
Quite true. I forgot to mention nitric acid. IRFNA is the usual form that it is used.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
Inhibited Red Fuming Nitric Acid (IRFNA) is probably a decent alternative to NTO; there is a small amount of Hydrogen Fluoride, about 2%, added for stability. This stuff is reasonable to handle, but only for those experienced in good laboratory practices. I've used the "one step down" from this Nitrogen Dioxide content in White Fuming Nitric Acid many times as a Nitration reagent in consort with Oleum, concentrated Sulfuric Acid containing excess Sulfur Trioxide.
Last edited by Oldfart1939 (2018-11-03 12:12:50)
Offline
Like button can go here
Here's a link to a YouTube presentation done by NASA in 1966 about NTO and Hydrazine:
https://youtu.be/bDRKeM9kKxs
Offline
Like button can go here
Thanks for that, Quaoar. I looked at that NASA Cearun page. Was it complicated entering in the chemical formulas for the propellants?
I like the idea of using very high expansion ratios for the nozzles when they are only intended to be used in space. You do have the problem of weight though for the high expansion ratios. Perhaps an aerospike could be used for the purpose? I was thinking the space shuttle underside ceramic tile material could be used for the aerospike because they are so lightweight. This would be lighter than metal or even graphite, but not as high temperature as graphite nozzles.
To deal with the fact the tile material is not as high temperature as graphite, could we use a combination of a bell nozzle with an aerospike beneath it? The bell nozzle at top would handle the exhaust coming out of the combustion chamber at very high temperature, but then below, as the exhaust cools, the aerospike would only have to encounter lowered temperatures.
This would have the advantage it would be altitude compensating so could even be used for ground launched stages.
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.”
Offline
Like button can go here
To answer Quaoar in post 7, yes, I think your Isp simulator software is doing a decent job, as far as it goes. I think you might find a nozzle with an area ratio = 1000 is not something easily built, and likely rather heavy.
...
GW
Quaoar, did you find the Cearun simulator did a good job for the sea level ISP? I’ve been using the free version of the Rocket Propusion Analysis program, http://propulsion-analysis.com/index.htm.
By comparing to known engines, I found it does a good job for the vacuum ISP, but poorly for the sea level value. I’d use Cearun eventhough it doesn’t have a good GUI, it it did do a good job simulating sea level Isp’s.
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.”
Offline
Like button can go here
If these codes have a nozzle efficiency model, use it at a kinetic energy efficiency of right at 98%, which is very typical of all physical nozzles, whether sea level or vacuum. That eliminates that error source. (Be aware that free-expansion designs operating far off design will have far lower nozzle kinetic efficiencies, too many folks assume high efficiency from such designs, and that is simply not correct.)
The engine cycle (how the pumps are driven) is the other big effect. Depending upon how that is done, not all the hot gas generated actually goes through the nozzle. You need to reduce both thrust and Isp by a factor representing how much mass does not go through the nozzle. If 2% of generated hot gas gets dumped overboard, reduce both thrust and Isp by a factor of 1.02.
The effect of reducing thrust reduces nozzle throat size to be consistent with the massflow actually going through the nozzle at the chamber pressure you select. This does not affect expansion ratio, but it scales your exit area down by the same ratio.
You have to do it that way, so that your total massflow drawn from the tanks is consistent with the mass ratio you use in your rocket equation estimates.
GW
Last edited by GW Johnson (2018-11-05 10:09:13)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
Does anyone here have any information regarding the coking residues of MMH and UDMH when reacted with NTO or IRFNA?
Offline
Like button can go here
Oldfart1939,
Maybe this will be of some utility:
COMBUSTION RESIDUES FROM N204-MMH MOTORS
Investigation of condensed and early stage gas phase hypergolic reactions
Contains some information not germane to your question, but useful to add to the discussion:
HYPERGOLIC PROPELLANTS: THE HANDLING HAZARDS AND LESSONS LEARNED FROM USE
Offline
Like button can go here
GW Johnson wrote:To answer Quaoar in post 7, yes, I think your Isp simulator software is doing a decent job, as far as it goes. I think you might find a nozzle with an area ratio = 1000 is not something easily built, and likely rather heavy.
...
GWQuaoar, did you find the Cearun simulator did a good job for the sea level ISP? I’ve been using the free version of the Rocket Propusion Analysis program, http://propulsion-analysis.com/index.htm.
By comparing to known engines, I found it does a good job for the vacuum ISP, but poorly for the sea level value. I’d use Cearun eventhough it doesn’t have a good GUI, it it did do a good job simulating sea level Isp’s.
Bob Clark
Hi, Bob
I've also tried to simulate the glorious F-1: the vacuum and sea level Isp of the real rocket were about 96% of the virtual one
Offline
Like button can go here
I (Be aware that free-expansion designs operating far off design will have far lower nozzle kinetic efficiencies, too many folks assume high efficiency from such designs, and that is simply not correct.)
GW
Please, forgot my ignorance: what is exactly a free expansion design?
Offline
Like button can go here
Quaoar,
Perhaps GW is referring to altitude compensating nozzles, such as aerospikes. If he is, then they're not generally (maybe never, since cooling a long spike is so difficult) as efficient as conventional nozzles designed to expand the exhaust product at a specific atmospheric pressure. The point is that the better altitude compensating designs maintain better kinetic efficiency across a much broader range of ambient pressures, offsetting or maybe even improving upon the overall efficiency of translating expansion of the combustion products into kinetic energy imparted to the vehicle the engine is attached to.
I would argue this from the point of what the ambient pressure is where the rocket burns most of its fuel. What's the efficiency at those pressures / altitudes? If your rocket only burns 20% of its fuel where the exhaust products are optimally expanded by the nozzle design, then what does that say about the nozzle's kinetic efficiency as it burns through the remaining 80% of its fuel?
There are other ways to correct the problem using conventional nozzle designs, such as extendible nozzles. GW has pointed that out to me a number of times, and of course, he's correct.
Offline
Like button can go here
Yes I was referring to spike-type altitude-compensating nozzles. At design, the streamlines produced by these are aligned about as close to axial as those by a conventional nozzle. Off design, they are not, being directed radially inward at high backpressure, and radially outward at low to vacuum backpressure.
Nozzle kinetic energy efficiency is essentially just a measure of how axial the streamlines are, which is constant at all backpressures in a conventional nozzle. eff = 0.5(1 + cosine alpha), where alpha is the half angle off-centerline of the nozzle streamtube boundary. Most conventional conical nozzles use a 15 degree half angle, which is 98.3% efficiency, hard to beat. For curved bells, you average the half angle at the throat and at the exit lip to get an average alpha, and plug that back into the same efficiency equation. The average with most designs is still quite close to 15 degrees.
For spike-type free-expansion designs, alpha is not fixed, and can take on quite large values. The advantage is that the effective expansion ratio is variable, so that the gas is expanded as much as possible, regardless of the backpressure. This zeroes the static pressure term in F = mV + (P - Pback)Ae. But as Pback goes to zero, the exit "lip" alpha goes close to 90 degrees! So your average alpha goes to near 45 degrees, or maybe a bit more. And so your nozzle efficiency drops below 85%.
This works to a finite backpressure, not true vacuum, for which the effective expansion ratio is infinite. You can get a higher Vexit, which is more effective than the backpressure thrust term, but at the cost of less axial collimation of your streamlines. There comes a low backpressure point (not zero !!!) where it is no longer worth the trouble of the more difficult design.
Free expansion designs are thus not a panacea for sea level vs vacuum operation, especially since it is usually necessary to actively liquid-cool the spike, thus making the design heavier than a conventional nozzle. The spikes get truncated precisely to avoid the extra weight, reducing the efficiency at really low backpressures, and also limiting the achievable expansion by limiting the effective area ratio to streamtube dimensions at that last point of contact.
Applied compressible fluid mechanics is a real bitch, ain't it?
GW
Last edited by GW Johnson (2018-11-08 13:03:24)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
Offline
Like button can go here
RGClark wrote:Quaoar, did you find the Cearun simulator did a good job for the sea level ISP? I’ve been using the free version of the Rocket Propusion Analysis program, http://propulsion-analysis.com/index.htm.
By comparing to known engines, I found it does a good job for the vacuum ISP, but poorly for the sea level value. I’d use Cearun eventhough it doesn’t have a good GUI, it it did do a good job simulating sea level Isp’s.
Bob Clark
Hi, Bob
I've also tried to simulate the glorious F-1: the vacuum and sea level Isp of the real rocket were about 96% of the virtual one
Was this among the standard propellants Cearun includes or did you have to calculate the elemental components of the propellants?
I think I tried the RPA program on the Atlas V engine and Delta IV engine and found it did poorly for the sea level Isp’s.
Bob Clark
Last edited by RGClark (2018-11-10 07:05:27)
Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):
“Anything worth doing is worth doing for a billion dollars.”
Offline
Like button can go here
Was this among the standard propellants Cearun includes or did you have to calculate the elemental components of the propellants?
I think I tried the RPA program on the Atlas V engine and Delta IV engine and found it did poorly for the sea level Isp’s.
Bob Clark
I've used standard cearun propellant: simulating the SpaceX Raptor (LOX-LCH4, 30 MPa, F/O 3.81, expansion ratio 40) I got: vacuum Isp: 369.6 s, sea level Isp: 348.7 s. Taking the 98% of these values we have: vacuum Isp: 362.2, sea level Isp 341.7: the vacuum Isp is near to the value of 360 given by SpaceX, the sea level is 11s higher than the 330s given by SpaceX.
Yes: you are right sea level Isp is less reliable: augmenting the throat area ratio the sea level Isp grows like vacuum Isp, without suffering of over-expansion.
Offline
Like button can go here