Talk:LGM-25C Titan II

Deployment Numbers

The deployment numbers cannot be correct. There were 54 silos at the 3 operational bases plus three training complexes with one missile each at Vandenberg temporarly on alert, so the maximum number of deployed missiles is 57 at any given time.Geomartin (talk) 08:45, 6 June 2009 (UTC)

Gov't Report

The text on the www.fas.org about the Titan II missile is not copyrighted material. It comes from page 233 of the U.S. Government report:

"To Defend and Deter: Legacy of the United States Cold War Missile Program" - 1996

by John C. Lonnquest and David F. Winkler USACERL Special Report 97/01 A study sponsored by the Department of Defense Legacy Resource Management Program Cold War Project 607 pages - illustrated

70.95-mb PDF format

The report is available at this URL.

http://www.cevp.com/docs/COLDWAR/1996-11-01952.pdf

Compare the Titan II section starting on page 233.

• The text on this current Titan II page comes from the US Air Force Titan II Fact Sheet at this URL:
• The link displayed is no longer validDwarner30uk (talk) 17:02, 4 October 2010 (UTC)

USAF Titan II Fact Sheet

Rusty 03:16, 19 Jul 2004 (UTC)

Titan II range

Titan II, as deployed, had a range of 5,500 nautical miles, not 9,000 miles. At one time the thought to deploy the Titan I Mark IV RV and warhead generated one flight that failed at staging and this is likely the source of the 9,000 mile range. The deployed Titan II had the W-53 and Mark VI RV, as correctly stated in the article.

Sources:

<"WS 107C, Titan II Weapon System Final Report, January 1965," held at the History Office, Peterson AFB, Colorado Springs,classified SECRET. The information cited is not classified.>/ <"Detailed Design Specifications for Model SM-68B Missile, Including Addendum for XSM-68B," held at History Office, Peterson AFB, unclassified.> 206.128.65.121 (talk) 02:11, 16 April 2009 (UTC)

No mention of fatal accident in Searcy Arkansas silo

| Source--Senor Freebie (talk) 11:50, 15 March 2012 (UTC)

Feel free to do so. Ckruschke (talk) 00:07, 16 March 2012 (UTC)Ckruschke

One lump or two

The LR-87 used in the first stage is a single engine with two nozzles. Andrewa (talk) 18:19, 24 November 2012 (UTC)

That is not correct, the LR-87-5 was a single chamber engine, which was fitted as a pair on the Titan II first stage. That is in contrast to the LR-87-3 of the Titan I, which was a two chamber engine with shared turbo-pump comparable to the modern RD-180.Geomartin (talk) 11:01, 25 November 2012 (UTC)

OHO! That would explain a lot. Sources? Andrewa (talk) 02:48, 13 July 2013 (UTC)

Found one that confirms this... http://www.astronautix.com/engines/lr875.htm Aerojet N2O4/Aerozine-50 rocket engine. 1096.8 kN. Out of Production. Isp=297s. Used on Titan 2 launch vehicle. Engines refurbished for space launcher versions from decommissioned missiles between 1974-1982. Configuration: twin fixed motors with individual turbo-pumped assemblies. Application: Titan 2 Stage 1. First Flown: 1962 ICBM. Sept. 1988 orbital. Dry Mass: 739 kg. Length: 3.13 m. Maximum Diameter: 1.14 m. Engine Cycle: Gas generator. Propellants: hypergolic nitrogen tetroxide and Aerozine-50, delivered at 750 kg/sec. Mixture Ratio: 1.93:1. Thrust: 1913 kN sea level. Isp: 259 sec at sea level. Expansion Ratio: 8:1. Combustion Chamber Pressure: 53.3 atm. Burn Time: 158 sec. Thrust (sl): 956.500 kN (215,030 lbf). Thrust (sl): 97,534 kgf. Engine: 739 kg (1,629 lb). Chamber Pressure: 54.00 bar. Area Ratio: 8. Thrust to Weight Ratio: 151.34. Oxidizer to Fuel Ratio: 1.93:1 Status: Out of Production. Height: 3.13 m (10.26 ft). Diameter: 1.14 m (3.74 ft). Thrust: 1,096.80 kN (246,570 lbf). Specific impulse: 297 s. Specific impulse sea level: 259 s. Burn time: 155 s.

Note however that the article on the engine currently [1] gives the weight of the two-nozzle LR-87-3 as 839 kg, while the weight of the LR-87-5 is given as 739 kg. As the LR-87-5 was lighter than the LR-87-3, the LR-87-5 figures are probably all for a pair of (single chamber) engines not a single engine... otherwise a pair of LS-87-5s would weigh 1478 kg, a lot heavier than the corresponding configuration using a single (two-chamber, assuming you're correct above) LR-87-3. Moreover that article lead states Though this powerful engine used two discrete combustion chambers, it is considered a single unit owing to both chambers using common turbomachinery, but then lists both the LR-87-3 and LR-87-5 as having Number of chambers: 1. And the lead to the section on the LR-57-5 reads A modified version for the Titan II used new propellants, nitrogen tetroxide and aerozine-50. The engine was generally lighter and simpler than its predecessor, partly due to the propellants being hypergolic (pyrophoric), which did not need an independent ignition system. So there is no mention of the introduction of separate turbopumps with the LR-87-5, which is the critical point in all of this.

So it's still a mess, with important omissions, contradictions and presumably, therefore inaccuracies (law of the excluded middle). But thanks for clarifying that... it's progress. Andrewa (talk) 03:26, 13 July 2013 (UTC)

Hi, getting difficult here. I was sure I red it in Stumpf's "Titan II - History of a cold war missile program", but the text about the engine is not very clear in this regard. So I guess I got the information from astronautix.com as well. However, Stumpf's book has a graph of the complete assembly showing two turbopump assemblies as well as two separate fuel and oxidiser lines each. I think on some pictures of the assembly avalable on the internet you can also see two exhausts stacks (on most, the angle is not good, there isn't a good closeup directly from behind. Anyway, I thought I go right to the source at Aerojet, but there aren't much details either. Then I went on to www.titan2icbm.org with that horrific mid90'2 webdesign, and they state The Stage I engine assembly was designated LR87-AJ-5. Figure 13 shows the major components of this engine assembly. The engine included two regeneratively cooled thrust chambers, two pump drive assemblies, and interconnecting lines and fittings, all supported by an engine frame. This website tells the same. So in conclusion, we have two thrust chambers and two pump assemblies mounted together on one frame, and the whole thing is called LR87-AJ5. Technically 2 separate engines, but assembled together in a way they are one. Now open to discussion if you call that 1 or 2. From my google search, the consensus on the web is 50:50.Geomartin (talk) 11:39, 13 July 2013 (UTC)

It does explain a lot. Am I right in guessing that single-nozzle applications of the LR-87-5 do not exist?

My suggestion is that we don't attempt to decide which is correct, but instead add a paragraph with good cited sources to the article on the engine itself saying (assuming that I'm interpretting what you say correctly) that some authorities regard the two-chamber assembly used on the Titan II, the GLV and so on as consisting of two engines, and some as a single engine. (And personally I think I can argue it both ways, but not without sliding into WP:OR in either case, so there's no point in trying to decide it anyway.)

We also need to clarify the specification tables associated with the LR-87-5 to say (assuming that I've got it right at last) that these figures are for a two-nozzle engine assembly. We can do this without needing to take any stand ourselves on how many engines this assembly represents. Or is that cheating? We don't want to be so clever with the wording that it becomes misleading, that defeats the whole purpose of Wikipedia.

We then need to do something about the infoboxes in the missile articles, and specifically, the number of engines lines. Will the template perhaps accept something like one two-nozzle assembly instead of just a number, linking to a footnote saying some authorities... and the footnote linking in turn to the more detailed discussion in the article on the LR-87?

Just suggestions at this stage. Thanks for coming back so promptly. Andrewa (talk) 11:35, 14 July 2013 (UTC)

Hi, I'd be happy with those changes. Let's call it 1 engine and explain in the text that LR87-AJ5 is an assembly consisting of two chambers with a turbopump each. That's what happens when 1950's engine designers don't take Wikipedia infoboxes into account. I think there should be a rule for engineers that their developments should be Wikipedia-conform. Probably an update of the LR87 article would be worth as well. From my memory it is the only engine family ever using all major fuel combinations: LOX/RP1, hypergolic, LOX/H2 and even LH2/F2. Would probably be worth to work on that thing a little. --Geomartin (talk) 12:23, 14 July 2013 (UTC)

Having spent two years working in software configuration management I feel for those who have to design and conform to spaceflight QA, let us not make it any more difficult for them. Then again, perhaps putting the doco through a Wikipedia consensus process would be an effective and efficient control... (;-> I'm half serious...

I'd still like to avoid taking a stand on whether it's one engine or two. Maybe it's not possible to sidestep the confusion in this way, but I'd like to have a go at it.

This is definitely one of the more significant, successful and interesting engine families, for many reasons. The success of Apollo impressed both the Soviet public and the workers at all levels in their aerospace industry, but they had little say in ending the cold war. The near-perfect record of the Titan II GLV was far, far scarier to those in authority in the then USSR, even if it lacked the propaganda potential. (And there was no Gemini 13!) But I'm not sure whether anyone else has published similar thoughts, and of course my WP:OR does not belong in the article. Andrewa (talk) 21:00, 16 July 2013 (UTC)

An interesting discussion. It never occurred to me to consider it a single engine. We always referred to them as "engines". To be precise (and having worked on them in the USAF, I know)...two thrust chambers and four turbo pumps (two per thrust chamber). That of course suggests that they are two engines, but they did share peripheral support systems. The hot gas generator (for pressurizing the propellant tanks) the lubrication system and the drive unit (a small reaction chamber where a negligible amount of propellant was reacted to drive the turbopumps for the thrust chambers) were critical parts of the engines. In that respect it is proper to say that since the thrust chambers could not operate independently, they must, therefore be considered a single unit.