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Misc

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This may be a stupid question, but can information be added to the article on whether the technical drawings and documents necessary to make more NK-33s still exist? From the paragraph "When the N-1 program was shut down, all work on the project was ordered destroyed. A bureaucrat instead took the engines, worth millions of dollars each, and stored them in a warehouse." it is not clear whether anything other than the physical engines themselves was preserved.

Well, after cancelling the N-1 poject all documentation and technological tools were supposedly abolished, only engines were saved by plant's stuff. Also authors should take in attention, NK-33 is modernised version of NK-15. Onboard N-1 NK-15 were installed. Hence, Russia should start producing NK-33 from very beginning. 84.204.75.2 (talk) 08:07, 19 August 2010 (UTC)[reply]

Thrust/weight

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The source says 136:1. An anonymous editor has twice replaced this with 120:1, on the grounds that back pressure reduces the thrust at sea level.

I disagree with, and have removed this, on two grounds:

  1. doing so is OR (it disagrees with the quoted source)
  2. vehicle performance is usually relatively modestly affected by sea level thrust/weight ratio, but more strongly affected by burnout mass (which typically occurs in near vacuum conditions) and the average Isp.

- (User) WolfKeeper (Talk) 14:57, 4 January 2008 (UTC)[reply]


- "sea level thrust/weight is far less relevant, since it is the burnout weight that mostly determines the mass fraction"

You don't even realize why this is off, do you?

N1 stage construction was ridiculously beefy- discarded N1 tanks currently survive as small buildings. The weight of engines- particularly these advanced ones- is much less of a factor.

In the more general case, a "classic" first stage is used to generate altitude, not velocity. The vehicle should not attempt drastic acceleration until reaching the upper atmosphere; this minimizes drag loss and aerodynamic heating. Thus, the relevant equation is not so much Tsiolkovsky's, but more Newton's second.

More modern first stages still have significant non-engine masses. First stages are loaded with reserve propellant. This covers performance shortfalls, as virtually all engines do not reach the point of mass production to guarantee high levels of quality. In addition, reserve propellant prevents a rough MECO, and high loads which may damage the payload. MECO is typically the first- or second-most violent design load. In earlier decades, lower-than-expected orbits and rough rides had to be tolerated. Now, especially for commercial flights but also for precision instruments on scientific missions, the launch provider is expected to guarantee a minimum altitude, and take at least reasonable steps to limit maximum stresses.

And then there's pressurant. First stages are often pressurized with nitrogen, not helium. As burnout altitude and speed are not that high, the weight of a tankful (or two) of nitrogen is higher, but considered acceptable. Meanwhile nitrogen reduces costs significantly. For a good-sized stage, the weight of pressurant gas at burnout can reach into four digits.

That's the difference between book knowledge and working knowledge. May 4, 2008 —Preceding unsigned comment added by 128.8.238.186 (talk) 19:41, 4 May 2008 (UTC)[reply]

The thrust:weight ratio of the engine, as with other metrics like power/weight ratio of engines is quoted at its peak. Since the peak is in vacuum, and since the source gives it as 136.66:1 you don't really have a leg to stand on. The weight and pressurant left in the ullage space of the vehicle is completely irrelevant to the isolated engine performance... which is what this article is about.- (User) WolfKeeper (Talk) 16:20, 5 July 2008 (UTC)[reply]

Wow! Here we have a truly classic wikipedia scenario: an anonymous editor who claims domain expertise from "working knowledge" of the subject -- but who (sadly) does not yet understand the vitally important role in wikipedia's culture of citing reliable sources -- is in apparent conflict with an experienced wikipedia editor. Of course the experienced wikipedia editor knows the article is only in compliance with policy when the assertions in the article match the claims made by the cited sources. The anon editor does not yet understand the "game" of finding sources to cite which match the truth. The predicted outcome? The experienced editor will win the conflict, but at a horrible cost: the domain expert will be disgusted with wikipedia and discontinue any efforts to contribute.

Wikipedia really cannot afford to spurn the attempts of any contributors, even those who do not (yet) play by the rules! So to the domain expert: Would you like help becoming a productive contributor? The welcoming committee is full of helpful people who would be glad to assist! Their welcome page is really quite good.... (sdsds - talk) 06:41, 6 July 2008 (UTC)[reply]

I just wanted to add that NK-33 isn't the engine with the highest thrust-to-weight ratio anymore, as SpaceX's Merlin 1D seems to have surpassed the 150:1 mark, how'd you guys feel about a correction? — Preceding unsigned comment added by 88.31.16.220 (talk) 00:08, 12 July 2012 (UTC)[reply]

Citation Requeste for Section

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The oxygen-rich technology lives on in the RD-170/-171 engines, and their RD-180 and recently developed RD-191 derivatives. These engines still use the multiple combustion chamber and nozzle topology like the RD-107/108 engines of Soyuz, thus preventing them from reaching the NK's high thrust-to-weight ratio. —Preceding unsigned comment added by 71.214.211.224 (talk) 21:27, 7 October 2010 (UTC)[reply]

I would like to see a citation for this, as the claim that the high thrust to weight ratio being due to the single combustion chamber versus multiple combustion chamber is dubious. The RD-171 class engines include a thrust gimbaling mechanism and thrust support structure. They also operate at higher combustion chamber pressures to achieve higher ISP's. These all would add weight to the engine, thus reducing the thrust-to-weight ratio. If the NK-33 included the gimballing, and thrust support, it's thrust-to-weight ratio would be somewhat lower as well. I believe that these are the main reasons for the lower thrust-to-weight ratios and not the claim about the combustion chamber topology as stated in the article. --71.214.211.224 (talk) 21:41, 7 October 2010 (UTC)[reply]

It's even more than that. The engines referenced use counter-rotating turbopumps in order to prevent torque from causing the rocket to spin, as the engines are designed to be operates in a single-engine configuration. The NK-33, on the other hand, uses a far simpler single-shaft design, with no capability to counter the turbopumps torque. The solution used on the Taurus II, on the Atlas V test firing using the AJ-26, and on the N-1 was to pair up engines, every engine would be paired directly against another engine which would be mounted backwards, in order words with it's turbopump spinning in the opposite direction. The two engines would throttle up and down together, countering each others torque. If the rockets needed to rotate, they would spin them slightly off-sync, in order to utilize that inherent torque, making rotation along the axis simpler.

The Soyuz 2-1v design, mind you has another solution, an additional motor for handling the steering, the RD-0110R, which has an additional flywheel to add torque to that motors turbopump, set in counter-rotation to the NK-33's turbopump.

Because these engines do not address the issue of turbopump torque, the natural tendency of any spinning item to turn it's base in the opposite direction to the spin, the NK-33 is far lighter than other engines, no use of gearboxes, or multiple turbopump control shafts, or any of the other techniques used by various engine manufacturers to prevent engine torque. It's nothing about topology, it's about simplicity. Those newer engines, as you pointed out, do a lot more than the NK-33, they gimbal, they include support structures, and they handle the torque their turbopumps put out internally. The NK-33 is an incredibly simple engine, and a very powerful engine. Trying to turn it into some kind of contest, just does not seem right to me.(Downix (talk) 06:56, 20 June 2011 (UTC))[reply]

Citations and sources are needed

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Please be sure that all additions to the NK-33 are verifiable. Any new items added to the article should have inline citations for each claim made. As a courtesy to editors who may have added unsourced claims previously, before Wikipedia citation policy is what it is today, many of the existing unsourced claims have been tagged {{citation needed}} to allow some time for sources to be added. N2e (talk) 22:33, 11 November 2010 (UTC)[reply]

User:The Bushranger has suggested in the edit comments that individual {{citation needed}} tags on specific claims be removed and replaced by a section tag indicating that more footnotes are needed. I'm afraid that won't quite do. I've seen many times that editors get up in arms if section content is deleted after a section footnotes needed (or article footnotes needed) tag has been on the article for many months. The general community practice seems to be that unsourced content should not be deleted without the specific claims, at least at the paragraph level, having been specifically and explicitly identified and left tagged for some time, generally a month or more. Thus, I will re-add the tags on the specific assertions I am challenging, since they have no source whatsoever. Cheers. N2e (talk) 22:43, 11 November 2010 (UTC)[reply]

Overuse of Citation, when the point referenced already has citations elsewhere in article. We're gaining redundant citations due to this overuse. I just had to remove a citation request because it was beyond silly to see a citation request for evidence that the engine had been tested in the US, when the article has a picture of the engine on a US test stand. I could, of course, added a citation, using the very article itself as a source, but that is getting silly. If you ask for a citation, you better check and make sure it's not already referenced elsewhere in the article. — Preceding unsigned comment added by Downix (talk • contribs) 18:23, 20 June 2011 (UTC)[reply]

AJ-26 is not exactly NK-33

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It is based on NK-33, but engine controls and some other things are changed. The main components- the thrust chamber, nozzle, turbopumps I believe are same as NK-33. --Aflafla1 (talk) 02:14, 3 July 2011 (UTC).[reply]

5 Valves have been changed. The thrust control valve and the mixture ratio valve. Then 3 more valves which formerly were pyrotechnic valves are replaced. But I can't find a open source for that. 77.3.171.155 (talk) 10:23, 11 January 2014 (UTC)[reply]

Merlin D has surpassed this

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Dear KlickitatGlacier this may be true but this does not mean the NK-33 engine is suddenly a "heavy" weight... It is still "extraordinarily lightweight" but it no longer has the crown of the "highest vacuum thrust-to-weight ratio of any comparable engine". Remember that these type of engines where designed in the late 60's early 70's and the NK-33 "only" lost the thrust-to-weight ratio crown in 2012! Not calling Staged combustion cycle (rocket) rocket engines "extraordinarily lightweight" anymore seems a bit harsh. If you want to shoot something "heavy" into space as "efficient" as possible in the year 2012 you probably need a Staged combustion cycle (rocket) rocket engine. --Jongfeli (talk) 11:58, 3 September 2012 (UTC)[reply]

So?
The article had an error and I corrected it, quit whining. --KlickitatGlacier (talk) 19:17, 3 September 2012 (UTC)[reply]
Well KlickitatGlacier I don know about the whining part, a simple "The article had an error and I corrected it" was enough for me. I was just trying to point out that the first part of the text you deleted (the resultant engines are extraordinarily lightweight) is about the Staged combustion cycle engine type and not specifically about the NK-33. So yes it has no longer "highest vacuum thrust-to-weight ratio of any comparable engine" but Staged combustion cycle rocket engines are still extraordinarily lightweight, thats all. Best Regards --Jongfeli (talk) 20:45, 3 September 2012 (UTC)[reply]
Yaaaawn, you really do like to hear yourself talk, don't you? --KlickitatGlacier (talk) 22:14, 3 September 2012 (UTC)[reply]
Well KlickitatGlacier if you say so. Looking at your history you sometimes seem a little obsessed about you fellow Wikipedians. I was not aware this was the "Talk about Jongfeli" Wiki, sorry. In the future I won't waist your precious time anymore. --Jongfeli (talk) 05:24, 4 September 2012 (UTC)[reply]

Subcooled?

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This article mentions "subcooled" in a couple of places. We have an article on subcooling, but I'm not sure if that is a proper explanation of the term as used here. --agr (talk) 13:19, 21 April 2013 (UTC)[reply]

I've removed the references to subcooling for now. subcooling would be the right place to link to in this context, but I found no sources that say subcooling was required for the engine to work (see Design section below). Subcooling was considered as an option to increase payload of the N-1 rocket. Zounds011 (talk) 19:40, 17 March 2016 (UTC)[reply]

Aerojet Merged Into Aerojet Rocketdyne

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Corporate history marches on, in 2013 Aerojet was merged into Aerojet Rocketdyne. – Conrad T. Pino (talk) 05:38, 31 October 2014 (UTC)[reply]

Design section

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This section currently contains this sentence: "In addition, since the NK-33 uses subcooled LOX and kerosene, which have similar densities, a single rotating shaft could be used for both turbopumps.[3] "

This has a few problems. 1. the referenced article does not claim anything about the turbopump. The reference is a useful source for the specifications box, but not much else. So it'd have to be removed here. 2. Subcooling kerosene and LOX have densities about 50% apart (0.8 and 1.2 kg/dm^3), "similar" isn't the word I'd use to describe that. Non-subcooled kerosene and LOX also differ by about 50% (see http://space.stackexchange.com/questions/14456/does-the-nk-33-engine-require-subcooled-kerosene-so-cold-that-it-turns-to-wax). The delta between subcooled and non-subcooled situations is small, I don't see it being a critical improvement. 3. Aerojet ran tests on an NK-33, running it on non-subcooled propellant and LOX for up to 140 s (http://lpre.de/resources/articles/AIAA-1998-3361.pdf). That indicates subcooling is not a prerequisite for having both pumps on a single shaft. 4. I can't find any source that says subcooling is a requirement to have a single-shaft turbopump.

There's also this sentence: "The turbopumps require subcooled liquid oxygen (LOX) to cool the bearings. " I haven't been able to find a source to verify or disprove that. The successful tests running on non-subcooled LOX suggest the engine does not need the LOX to be subcooled.

So what I have is 2 statements that aren't correct, a document that confirms they aren't correct, but no sources that I can refer to to make correct statements on the design of the turbopump. For now, I've removed the incorrect statements. Will look for more info, but haven't found anything relevant yet. Zounds011 (talk) 19:37, 17 March 2016 (UTC)[reply]

I've added Mr Eberly's (Orbital ATK Antares Deputy Program Manager) comment on Antares's ground support facilities as another citation for NK-33's subcooling. PSR B1937+21 (talk) 12:56, 18 March 2016 (UTC)[reply]

Further, the section seems to state that RD-180s were manufactured in the US. I don't think the technology has been used yet in any US-manufactured engine; the ones coming from Aerojet were refurbished NK-33s of Soviet manufacture. — DAGwyn 71.179.2.197 (talk) 07:34, 8 September 2017 (UTC)[reply]

Schematic layout of the NK-33 (Pre Burner.png)

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Hello Beesmill. I appreciate the time you took to create a schematic layout of the NK-33 but unfortunately it is not correct. The NK-33 is a oxygen-rich staged combustion cycle engine, the schematics shows a fuel-rich engine like the RS-25. Regards, --Felipe (talk) 05:46, 19 September 2016 (UTC)[reply]

BeesMill here: Jongfeli, When you get a change, please take a look at the film, Cosmodrome 2008 49mins. The Soviet Engineers also developed the Jet Engines for the Strategic Long Range Bombers, therefor efficiency was paramount, thus rocket (liquid propellant engines) did not pose any more challenges than aircraft engines. In 1962 the basic design with a Liquid 02 and Kerosene are the mixtures are pumped into the combustion chamber, where they are ignited, pushing the rocket forward. The higher the pressure in the chamber, the higher the performance of the engine. The high pressures are driven by pumps, driven by gas turbines fired in the Pre-Burner, a mini combustion chamber that draws off part of the 02 and fuel supply, The time on the N1 was running out to beat the USA, so a solution was needed. Only small but efficient changes were feasible. An open cycle was used by both USA and Soviets where exhaust products from the pre-burner are dumped as exhaust away from the main combustion chamber making the rocket inefficient. The Soviets needed to CLOSE the cycle, a risky step. The exhaust from the pre-burner was channeled back into the combustion chamber. This was a very explosive situation but could boost the lifting process, a boost of 25%. Welcome to the closed cycle NK-33. — Preceding unsigned comment added by Beesmill (talk • contribs) 07:03, 19 September 2016 (UTC)[reply]

Hello Beesmill, yes I saw the movie. a couple of times actually. In Pre Burner.png on the NK-33 page you see a fuel-rich pre burner where ALL the fuel goes trough the pre-burner. However the NK-33 is a oxygen-rich engine, this means that ALL the oxygen goes trough the pre-burner and not the fuel. Please have a look on the RD-180 page, it to has a schematic overview of the engine, the NK-33 basically works on the same principle. Regards, --Felipe (talk) 08:45, 19 September 2016 (UTC)[reply]

I now understand. The film "dumbed it down" meaning the schematic is very simplified. Thanks for the clear visual, I now need to review the documentsBeesmill (talk) 08:55, 19 September 2016 (UTC), again thanks. Beesmill[reply]

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