I keep meaning to write something about US nuclear posture soon, but the day-job is all consuming right now.
In the meantime, I want to point out the fascinating debate begun by Geoff Forden on whether the Safir, Iran’s new space launch vehicle, has two-stages — as Iran claims — or uses a small third stage to place the satellite in orbit. As usual, the Arms Control Wonk readership is terrifyingly astute and engaged. (See: Why Would a 2 Stage Safir Be Surprising? and the oddly-titled, Congratulations, Iran).
For those of you not following the debate in the comments, using a two-stage rocket to place even a small satellite in orbit would signal a significant advance beyond the improved Scud technology embodied in the North Korean Nodong/Taepodong families.
David Wright, Co-Director of the Security Program at the Union of Concerned Scientists, sends along a preliminary existence proof that suggests a two-stage missile based on N2O4/UDMH fuel (similar to China’s mid-1970s ballistic missiles) would be sufficient to place a small satellite in orbit with only two stages:
Based on these numbers, I consider the following model:
Stage 1:
Propellant mass: 20 t
Fuel mass fraction: 0.91
Isp(SL) = 258 s
Isp(vac) = 280 s
Burntime = 140 s
Thrust = 361.2 kNStage 2:
Propellant mass: 4 t
Fuel mass fraction: 0.85
Isp(vac) = 298 s
Burntime = 140 s
Thrust = 83.4 kNSatellite mass = 27 kg
Payload shroud = 50 kg, released at 150s when the second stage is above 100 km.I assume a round earth, and use a model atmosphere to calculate drag during launch.
Using the parameters given above, I calculate a burnout speed of 7.6 km/s at an altitude of about 240 km, which is consistent with the Omid launch.
Moreover, this trajectory gives a first-stage burnout at 66 km altitude, which appears to be consistent with a report from Iranian radio about the launch.
I should caution that I do not have enough information about the Safir launcher to suggest that these represent the actual values for the booster, or that it uses N2O4/UDMH propellant. My goal in presenting this calculation is to show that booster parameters consistent with those that were developed several decades ago would allow a two-stage launcher of the size of the Safir to place a small satellite in orbit, without the need of an additional kick stage.
You can read David’s entire paper here. (To the four eager beavers who downloaded it before 12:48, I’ve uploaded a slightly edited version.)
To be clear, as David is in his introduction, the analysis is not intended — and does not demonstrate — that China has assisted Iran or that the Safir is an Long March rocket:
This analysis is not intended to suggest that China has assisted Iran with its development program or transferred technology to Iran, since this kind of technology has been around for decades. The analysis uses Chinese launchers as a concrete example of what booster parameters have been achieved.
Blog Posts
Get E-mail Updates
A kerosene fuel with N204 in excess of 50% will also allow a two stage rocket to put up Omid.
Also it seems that the Safir is not designed to ever have a substantial third stage. It’s liftoff thrust is at the lower end of what it needs as things stand now. Any more weight on that stick will just turn the thing into a pyrotechnic paper weight. If however the Safir is designed for strapon 0 stage solid motors, such as 2, 3 or 4 sejil first stage motors, then the Safir, with or without a third stage, could lift a more substantial payload to leo and perhaps beyond.
— Azr@el · Feb 11, 01:59 PM ·
The “more than 7000 meters per second speed at which the satellite was deployed” – quoting President Ahmadinejad yesterday – is also consistent with David’s computations.
— Paul · Feb 11, 05:59 PM ·
Well done, [Dr.] Wright!
As your calculation example clearly shows, a two-stage-to-orbit missile is feasible – even with storable propellants and such a small launcher (minor points of critique: burnout-acceleration for the upper stage of ~12g is a bit high; slightly too high masses; slightly too high liftoff-acceleration; and why does it ALWAYS have to be UDMH, although we have, due to a visible tank-volume-ratio of ~2:1, explicit indication against that on the Safir-1.stage?).
I’d say that at this time all available evidence points into the direction of the Safir being only two-staged (especially the existence of a second orbital object with a distinctively lower apogee than the other – as Geoff already demonstrated, the necessary delta-v-difference between both objects is in the range of what a simple spring mechanism can produce).
Of course, this implies lower structural-mass-ratios as on the Scud- or the NoDong-line of missiles.
But, in my opinion, not necessarily a radically new kind of construction technique or completely different materials (aka a GREAT technological breakthtough)!
I think the necessary reduction in dry-weight could have been achieved by simply using 1.5mm-steel-sheet-metal (like on Scud-C) instead of 2.0mm-steel-sheet-metal (like on the Scud-B and presumably also the NoDong-A and -B) as a construction material for the tanks in combination with the application of smaller fins from the Scud-line of missiles on a stretched NoDong-airframe (this has already been accomplished on the Ghadr-1/NoDong-B), albeit at an accompaning reduction in structural resilience (but that would probably be acceptable for a space launcher!). And it should also not be forgotten that thanks to the mass distribution generated by the fully-fueled upper-stage, it was possible to omit the separating bulkhead between both IRFNA-tanks (like on the Kosmos-B1, which had a similarly modified R-12/SS-4/Sandal as the 1.stage) on the lower stage (in contrast to the Ghadr-1/NoDong-B).
And we have to once and for all bid farewell to the unwarranted predjudice that the Scud was “nothing but a primitive V2-clone” (in certain aspects like e.g. the use of unnecessary stringers or operational safety, the Scud was even superior to the Titan-II!).
So, in essence, the Safir IRILV might probably not be the end of “NoDongism”, but the final culmination point of that technology, even though this came at the price of lower margins of safety and distinct operational constraints.
— Jochen Schischka · Feb 12, 11:53 AM ·
According to “Paul” and “Pedram” in an earlier thread, President Ahmadinejad stated that the stage 1 and stage 2 burn times were 3 and 5 minutes respectively.
While he may make some objectionable comments in other areas, President Ahmadinejad’s statements regarding Iran’s technical progress have generally turned out to be correct (witness the various claims for numbers of operating centrifuges, which were subsequently confirmed by IAEA inspectors, or the claim that a previous rocket had reached ca. 200 km altitude, which was shown to be at least plausible on this blog).
Additionally, there’s every chance that the launch was tracked by western intelligence agencies, who surely would be capable of determing the burn times. So deceit in this direction would serve no purpose other than to leave him open to later exposure and loss of face.
I appreciate that 40 seconds can be a big discrepancy in rocketry, but after allowing for the fact that he was probably rounding his figures, I think that President Ahmadinejad’s statement can be considered as “evidence” – evidence that tends to confirm David Wright’s proof of plausibility.
It might even be worth taking President Ahmadinejad’s statement at face value, and seeing what the assumption of a 180 second first-stage burn time would imply.
— Hairs · Feb 12, 02:33 PM ·
Jochen:
According to www.gwu.edu/~nsarchiv/NSAEBB/NSAEBB39/document1.pdf a Scud casing is stainless steel (somewhere else I saw 30CrMoV9 as the material) with a 142 ksi strength (= ca. 979 MPa). So a change to maraging steel (e.g. the 1800 MPa maraging steel that the Indians have been working with), could easily give your required thickness reduction without any loss of structural resilience.
Such a change might be a departure from historical Scud / Nodong materials, but it shouldn’t be such a big jump as, say, designing a completely new engine.
— Hairs · Feb 12, 03:12 PM ·
The longer burn time would imply that the Safir had to contend with more gravity drag than David Wirght’s analysis allows for. Also note that the Safir could have gone up without UDMH as fuel, but it would have needed a much more energetic oxidizer than 27% N2O4 to get that sort of performance out of a kerosene based fuel.
Furthermore upon reading Iranian goals for the Safir, it becomes more obvious that the Safir is designed to put up payloads of close to 100 kg. And potentially more with the addition of a very small third stage kick engine. The Safir appears to be at it’s limit in terms of mass able to be lofted by it’s 1st stage, so unless there are plans to outfit it would solid motors strap on boosters, it’s fair to say the Safir could not contend with a more robust third stage. But should it be so equipped, 540-870kg to leo would not be out of the question.
— Azr@el · Feb 12, 03:16 PM ·
Ahmadinejad also mentioned in his speech that the satellite was put into orbit at “speeds exceeding 7000 m/s”.
— Paul · Feb 12, 05:44 PM ·
Hairs:
I think it is generally advisable to use a healthy dose of skepticism toward the statements of politicians. They could have misunderstood things. They could have been misinformed by subordinates. They could have been misquoted by the media. They could intentionally overstate (consider for example the famous claim of Mr. Chrushchev about “producing missiles like sausages”). And there is of course always the possibility of deliberate maskirovka (e.g. Walter Ulbricht’s statement about “nobody having the intention of constructing a wall” – a mere two months before the Berlin wall actually WAS built according to plans and preparations considerably older than that declaration).
But if said statements DO agree with verifiable facts, there is no reason to not believe them.
As you might have noticed, i AM willing to believe in a two-staged configuration or a spring mechanism for payload-separation – simply because that does agree quite well with the (up to now) known facts. The ~300sec-figure for upper-stage-operation could also be credible (if you assume a modified R-27/Zyb/SSN-6/Serb-vernier-engine with a dm/dt of ~12kg/sec), but since i consider this issue yet to be neither well-established nor verifiable (nor falsifyable) with absolute certainty, i’d suggest to handle that number with caution.
On the other hand, the 3 minute-figure for lower-stage-operation does raise certain questions. The net-liftoff-acceleration of the Safir is ~0.3g according to measurements. Based on photographic evidence (and the evident general design-heritage of that missile), i tend to believe in a 1.stage-propellant-combination of IRFNA/Kerosene (which yields an Isp(SL, eff) of ~224sec for a chamber-pressure of ~55bar and ~226sec for ~70bar according to simulation/calculation/estimation). If we accept a figure of 26t for liftoff-mass, then the sea-level-thrust must range around 33-34t to explain that acceleration. Such a thrust (associated with ~70bar) would require a dm/dt of about 150kg/sec, leading to an overall fuel consumption of 27t – one ton MORE than the complete missile! Now let’s work through that example again with a thrust level of only 27t (connected to 55bar p-c; verifyable by analysis of photographic evidence of all shots of Shahab-3/Ghauri/NoDong-A and Ghadr-1/NoDong-B to date): This time, we get a dm/dt of ~120kg/sec and a total fuel consumption of ~21.7t. But since the liftoff-mass this time can only be ~20.8t, this would still be highly inconsistent! The real figure for total burn-time would in any case be closer to ~130sec – and that is definitely more like TWO minutes even by non-engineering standards.
But i wouldn’t put too much emphasis on that little discrepancy – this could simply be a slip of the tongue, a misunderstanding or even a translation mistake.
BTW, there is a plethora of examples of classical deception coming from Iran (e.g. according to the Teheran September-parade of last year, the “Sedjeel” solid-fueled missile would be nothing else than a repainted Hawk-SAM) – the real art is to filter out the “believable lies”…
— Jochen Schischka · Feb 12, 06:30 PM ·
To Hairs:
But higher-alloyed steel would also be much more difficult to weld and machine!
The 1.5mm-material, on the other hand, could be processed on the same kind of tooling/machinery. What is more, the thinner sheet-metal would already be familiar to the Iranians and/or North Koreans from the Shahab-2/Scud-C. And probably available from the same source.
I agree with you on the (first-stage) engine, at least to a certain extent. A completely new engine, built for other propellants than that of the NoDong would be extremely difficult to indigenously design (this probably does not include modifying an existing design…) without any experience on that sector.
To Azr@el:
Wow, you’re a real fan of NTO, aren’t you?
I personally think you somewhat overestimate the effect this would have on the first-stage performance while (haven’t i written that somewhere before? sounds familiar…) COMPLETELY neglecting the visible tank-volume-ratio (a little tip: look out for the typical double weld-seams characteristic of tank-end-domes).
But if you have striking evidence to back up your theory, you are of course highly welcome to share that with us!
— Jochen Schischka · Feb 12, 07:00 PM ·
There is no disinformation/deception about Sedjeel for those who follow Iranian military developments. The Sedjeel project at the Iranian Air Force involves retrofitting a Hawk surface to air missile onto F-14 fighter jets. This is completely distinct, older, and separate from the Sedjeel solid fuel ballistic missile.
— Paul · Feb 12, 07:01 PM ·
Iranian naming conventions again! I used to believe the Iranians were conducting a campaign of mass disinformation by giving the same name to different weapons systems; for instance “Sejil” has been used as the name of an absurdly long ranged I-Hawk missile modified into an BVR AAM fielded by Iranian F-14s in 86’ to replace their exhausted stocks of Phoenix Aim54s , the name of a hand tossed shaped charge grenade , the name of a discontinued unguided rocket barrage system deployed by Guards and yes it is also the name of their latest two stage solid rocket IRBM based on PRC tech transfer.
In reality there is no disinformation campaign, the confusion arises as a product of Iranian decentralization of their weapons program and the religious overtones of Iranian culture. The Japanese had an equally confusing naming protocol during the Pacific War, drove U.S. intel batty, but made perfect sense to the Japanese…I’m sure the same applies to the Iranians.
— Azr@el · Feb 12, 07:53 PM ·
Some afterthoughts:
Maybe the 3min-figure includes a coast-phase of ~50sec (the type of staging technique could hint at something like that, and this would additionally be consistent with an orbit-optimized trajectory).
Expanding on my thoughts considering Azr@el’s suggestions about a higher NTO-share for the oxidiser, i must say that empirically, NTO does not per se allow MUCH better Isp-values. What is more, as NTO now goes along with more fuel instead of oxidiser in the same overall volume due to a different optimum-O/F (and oxidiser density), that would decrease the overall propellant mass somewhat (due to the lower average density of that propellant combination) while changing nothing in the dry-weight of the associated tankage, so all in all the gain is less than would be expected by only looking at the Isp-values (even if these would be considerably higher) – especially in a lower-stage application (where specific impulse is not the all-out dominating factor).
In retrospect, maybe Hairs and i are both partially right: maybe the (hypothetical) 1.5mm-sheet-metal IS somewhat higher-strengthed material – but still weld- and machineable with the original tooling; By this, the risk could be shared between a moderately higher strain on the production-chain AND a minor decrease in safety factor on the missile (this would of course also be valid in regard to the Scud-C). Has anybody access to conclusive (declassified) data on the exact type of sheet-metal material for Scud-C seized in the Kuwolsan-incident in 1999?
— Jochen Schischka · Feb 13, 08:33 AM ·
Paul, Azr@el:
Let me give you some other examples of iranian deception. Have you noticed that from time to time, parts of the tyres of the Shahab-3-MEL were painted in the same colour as the felloes (whose diameter gives a nice scaling-tool for photo-measurements)? What about the Shahab-3/Ghadr-1-hybrids (either Shahab-3/NoDong-A with the conical guidance-section and triconic warhead or the large fins on a Ghadr-1/NoDong-B) paraded in 2006 and 2008? And what about the “Azarakhsh” displayed in 2008, that was nothing else than standard F-5Es and Fs painted in the colours (including the writing “Azarakhsh” on the tail) of the REAL Azarakhsh-prototype (different wing and air intake than a F-5E)? Or the not very expertly photo-shopped image of the Zelzal-test last year?
Let’s be honest, they’d be stupid if they would NOT try to cause some (additional) confusion about their newest and most secret weapons systems.
Of course, especially considering the “names-game”, the misunderstanding can also happen on the receiving end – all the more is it essential to retrace the facts and come to a better understanding of the whole matter by means of plausible reconstruction instead of simply parroting the “official” (mostly translated and/or transcribed by not-too-reliable press agencies, be that iranian OR western ones) statements unchecked.
And now, gentlemen, i’d suggest to concentrate more on the technical issues of the Safir, since that is clearly a much more worthwhile topic, isn’t it?
— Jochen Schischka · Feb 13, 11:47 AM ·
Jochen:
Maraging steels are easily machinable (because of their low carbon content) and certainly have better weldability than the stainless steel of a Scud casing – particularly if that Scud casing is 30% Cr! Prior to aging maraging steels are actually quite malleable too, and you can cold roll them 70% without any problems. So it would be possible to form a casing to its final dimension comparatively easily, and then treat the steel to get the final hardness.
The biggest problem I’m aware of in welding maraging steels is that you tend to lose material toughness in the heat affected zone. This can be largely alleviated by post weld heat treatment, but it becomes progressively more difficult to do with higher numbers of passes and / or thicker materials. However, something as thin as 1.5 mm really shouldn’t present any great problems.
Back in 1998 Mr. Manzarpour was jailed for supplying maraging steel to Iran, and there has been no shortage of stories about Iran’s efforts to import maraging steel since then. There’s also plenty of suspicion about Iran’s research efforts related to maraging steel. So all in all, if they’re not already using maraging steel then I’m sure they’re working on it.
— Hairs · Feb 13, 12:34 PM ·
You want the shortest coast phase possible after first stage separation, in order to maximize performance bye expending the fuel as close to the center of the earth as possible. Letting the second stage rise to the top of its trajectory before starting the engine is a characteristic of spin-stabilized solid propellant upper stages, where the stage must be pointed parallel to the ground before firing.
— Murray Anderson · Feb 13, 02:37 PM ·
See this photo
The label in Persian reads “2nd stage engine”.
— Paul · Feb 13, 03:42 PM ·
My only point with N2O4, which yes does yield a higher ISP as it’s % increases with respect to Nitric Acid as can be seen from an ISP chart of the AK20 series, is that there are alternative higher ISP oxidizers the Iranians may have turned to as opposed to AK20I.
As far as the Safir, a cursory review of the video yields; A) A discharge which is most likely the pyrotechnic start up charge for the turbine, B) The squeal of the turbine with a small red burn of the hypergolic activation of the chamber with tonka and the oxidizer C) A large explosion as the kerosene based fuel final hits the chamber. The burn is yellowish with a tinge of red. A review of the daytime launch of the Kavoshgar-1, which gives a very good view of it’s MEL, shows almost exact characteristics from which it could be safe to infer that launch was a test of Safir’s first stage. Which allows us to observe more clearly that the exhaust is greyish white which is not the case with Shahab’s and NoDongs. ( On a side note, something small and burning is being ejected from near the base of both at about 150m)
From the large aft flame of the rocket it seems apparent that the Safir is burning a hydrocarbon fuel that is not fully undergoing combustion in the chamber. So this not an efficient or uprated engine design, this is more likely a rather nominal engine design at the lower end of the potential ISP curve. But it has far less of the brown sooty exhaust of a scud b, Shahab 3 or NoDong launch…highly suggesting that they’ve swapped out the oxidizer for something more energetic that is causing a higher chamber temperature => faster species velocity and thus less turbulence in the nozzle => less generation of soot. What that could be is speculative, my guess? they’ve increased the ratio of NTO to Nitric acid.
— Azr@el · Feb 13, 03:55 PM ·
A few more photos:
Photo 1
Photo 2
(second stage engine)
Photo 3
(second stage engine)
Photo 4
Photo 5
Photo 6
Photo 7
— Paul · Feb 13, 06:31 PM ·
Hairs:
Have you ever wondered about WHY the Scud uses stainless steel for the tanks (BTW, the fins and the airframe of the tail are made from aluminum-alloy)?
The high Chromium-content (i’m investigating at the moment if this really is 30CrMoV9, but haven’t found anything conclusive yet) is required for increased resilience against one of the propellants: nitric acid.
So what we’d need is basically a weldable (preferable by means as simple as possible), cold-rollable, nitric-acid-resistent steel with ~20-40% higher strength, but not considerably higher density than that of the Scud-B. Oh, and the weld-seams should better be resistent to IRFNA as well…
(I think the Russians maybe used MAG-welding on the Scud-B, but this is more predjudice than knowledge – can somebody with higher insight into welding of stainless steel than me check this issue?)
— Jochen Schischka · Feb 13, 07:04 PM ·
Murray Anderson:
As far as i understand, in this case the upper stage was more or less optimized to provide most of the “circular” part of the delta-v (boosting mostly parallel to earth while still climbing with the inertia from lower-stage-operation), while the lower stage was mostly responsible for the maximum orbit height; And for the burnout-speed being adequate for orbit at apogee (or rather perigee), the intended cutoff has to be timed to coincide approximately with the greatest achieved height, thus perhaps a coast phase between lower stage cutoff and upper stage ignition (if the dm/dt could not be adjusted during design without spoiling the best attainable Isp -> bought engine optimized for another missile?). But of course, this is only a theory.
— Jochen Schischka · Feb 13, 07:29 PM ·
Paul:
Wow, this picture (obviously from the February-2008-display) shows us the backside of the upper stage engine in high-resolution (although i think this is only a testrig for ground-tests because of some peculiarities of the plumbing and the turbopump-canister).
And there are two additional nice pictures of the simultaneously displayed Safir-mock-up (large fins, illogically short cable duct) on that page:
http://www.farsnews.com/plarg.php?nn=M430985.jpg
http://www.farsnews.com/plarg.php?nn=M430917.jpg
— Jochen Schischka · Feb 13, 07:45 PM ·
Azr@el:
I hope you are aware of the fact that supplanting propellants requires at least a completely different turbopump (not to speak AGAIN of a different tank-volume-ratio…) due to different densities and optimum-O/F-ratios (otherwise you waste A LOT of potential Isp)?
And that building chambers long enough for a complete combustion inside of the chamber is highly uneconomical (and thermodynamically problematic), especially in respect to Isp?
I agree on points A), B) and C), but i can’t follow you on most of the rest. I can find no “brown sooty exhaust”, since both launches of the Safir occured at night – and thus any detail of interest of the rocket exhaust vanished behind a bright white-yellow blob in the official footing. On the Kavoshgar-1, i can deduce a thrust-level of about 25-27t (due to a visible length of the afterglow of the rocket-exhaust of approximately 16 meters) being suggestive of a standard NoDong-engine (in case of an uprated “NoDong+”-engine with 33-34t, it should be at least ~18m). And i have difficulties perceiving the “greyish white” colour that you mention (BTW, the quality of colours in the videos i have seen of the Kavoshgar-1-launch generally is not the best due to backlighting conditions)…
— Jochen Schischka · Feb 13, 08:50 PM ·
Jochen
I see your point, but a coast period implies an greatly over-sized upper stage engine, which would make it harder to get the rocket in orbit. If the thrust of the engine is equal to the weight of the stage propellant, then you’d get 280-290 seconds burn time, under reasonable assumptions about specific impulse.
You’d need an engine with 2 or 3 times the thrust of upper stage weight to require a coast period, I think.
— Murray Anderson · Feb 14, 12:14 AM ·
Murray:
Exactly the other way around!
The thrust of the engine doesn’t have to be able to lift (or positively accelerate) the fully fueled upper stage against the gravity-force in my scenario – that would, at least to a certain extent, already have been accomplished by the lower stage. Since there is no noteworthy atmosphere at the altitudes involved (another potential reason for a coast-phase), flight at extreme angles of attack (boosting parallel to the surface of earth while the overall vector of the trajectory would still point upwards -> no gravity-losses!) wouldn’t have any negative effects anymore.
(BTW, there are many examples of such lower-thrust-than-weight upper stages on existing LEO launchers NOT derived from ballistic missiles, e.g. the EPS L10 and ESC-A of the Ariane-5 or the Saturn-IB’s S-IVB, just to name some; This is exactly my point – we have to stop thinking in “ballistic missile” categories).
— Jochen Schischka · Feb 14, 12:19 PM ·
I’ve been working on establishing engineering specifications for the Safir launch vehicle. I’ve started by looking at the first stage engine.
I’ve used the following three images as references:
Image1 Image2 Image3
Image 1 is the presidential inspection photo of the base of a Safir rocket.
Image 2 is an image of a Safir first stage engine likely similar to the one used in the orbital launch and which is presumed to be demonstrative with few significant dimensional modifications.
Image 3 is an image of a cross section of a Scud engine to estimate combustion wall thickness.
I’ve done an analysis with error propagation to determine the Nozzle Expansion Ratio:
Ratio Analysis
I presumed that the diameter of the rocket is known as 1.25 meters +/- 3 millimeters (WAG01). I presumed I could read a feature in an image with an error of +/- 1 pixel.
I had to make a WAG about the thickness of the combustion chamber (WAG02) by estimating the thickness of the scud combustion chamber from Image 3. I’m presuming that the construction techniques used in the Scud are representative of those used in the Safir engine (WAG03).
I’d welcome any comment on the numbers used (yes, I know I violated significant digit conventions). I’d like any comments on whether any of the numbers used are terribly wrong. Have I established that the expansion ratio is about 11.4 +/- 0.6 given the above information? What would I have to change to get a better answer?
Thanks…
— Ed LeBouthillier · Feb 14, 12:37 PM ·
Jochen, you wrote:
“Have you ever wondered about WHY the Scud uses stainless steel for the tanks (BTW, the fins and the airframe of the tail are made from aluminum-alloy)?
The high Chromium-content (i’m investigating at the moment if this really is 30CrMoV9, but haven’t found anything conclusive yet) is required for increased resilience against one of the propellants: nitric acid.”
I would guess that the main reason stainless steel was used for Scud casings was simply because high nickel maraging steel wasn’t around when the Scuds were first designed. Additionally, for a missile that was to be produced in the hundreds, or thousands, it probably wasn’t worth the expense of going to better materials in later decades because the additional range wasn’t worth the money required (much better spend the money on a new engine / rocket design).
Yes, the high Cr content in a Scud tank provides protection from nitric acid, but only by passivation to a surface layer of oxide. In contrast, the high nickel content of maraging steels means that they can take advantage of the fact that NiF2 is extremely stable under acidic conditions. For example, highly-aggressive pure HF can be carried in tanks made of Monel – something you couldn’t do with a chromium-rich stainless steel. Similarly, ca. 1% HF in your nitric acid will form NiF2 on the inside of the fuel tank, protecting it from further attack – it’s where the “inhibited” comes from in IRFNA.
Maraging steels can be cold rolled to final dimension before hardening, offer better machineability and weldability, and can be made to withstand acid attack better than high-chromium stainless steels. Their biggest disadvantage is they are difficult to make / procure, and they are much more expensive than most stainless steels or aluminium.
I’m not claiming that the Iranians are using maraging steels, only that if they did use them they would be able to reduce their launcher’s weight without any loss of structural integrity. As such I think it is a good bet that if they’re not using maraging steels already then they are working on it.
— Hairs · Feb 15, 12:40 AM ·
A clear picture of the second stage engine:
http://www.fresh.co.il/vBulletin/showpost.php?p=3245786&postcount=1
— Tal Inbar · Feb 15, 09:50 AM ·
To Tal Inbar:
I doubt that this is the engine used on the Safir second stage (which obviously is, in contrast, a design with two gimbaled chambers fed by a common open-cycle turbopump in a canister submerged inside of the lower propellant-tank – all features of the R-27/Zyb/SSN-6/Serb-vernier-engine, which also has, by coincidence, rather similar, if not identical, dimensions…).
The engine on the picture provided by you looks exactly like an Isayev S2.711 from a Sayad-1/V-750V/11D/Desna/SA-2b/Guideline-B (or mod.1) surface-to air-missile to me (IRFNA/Tonka, open-cycle, 3.1t thrust-SL; fixed to the structure of the missile by the cross-like structure at the head of the chamber). You might notice that the belonging missile is also on display at the same exhibition (at present in Teheran? Can anybody please provide more information on that exhibition?). BTW, the SA-2 is a two-staged system, too: double-base booster + liquid sustainer (this could explain the little misunderstanding).
— Jochen Schischka · Feb 15, 12:55 PM ·
To Ed LeBouthillier:
Once again, you impress me with your professional modus operandi!
Your numbers do agree quite well with my own measurements/estimations, although i think that you overestimated the wall thickness by a factor of ~2, so the real number would be more in the order of 5-6mm (look for example how extremely thin-walled the Scud-engine in Image3 is for comparison). I got an (inner) nozzle end diameter of ~620mm (photo-measurement) and i estimate an inner throat diameter of ~200mm (backed up by data from a cross-sectional drawing of an obviously never produced very similar thrust-chamber in an old soviet textbook about rocket engines) and thus an area ratio of about 9.6 (consistent with a ground-start engine – i consider a value of over 11 somewhat high for the chamber pressures presumably involved). The corresponding values for the 9D21 of the Scud would be (source: “Dienstvorschrift DV 11/22” of the former east german army): d(t)=124mm; d(e)=400mm; A(e)/A(t)=10.4;
There are also clear indications for an Isayev-heritage of this engine: look at the triangular three-fold pneumatics-coupling (marked with “PG” on the Scud) and the two boxes to the left and the right of the nozzle (marked with “Sh37” and “Sh38” on the Scud) in Image1. All the hatches on the engine compartment are in the same location (only scaled up accordingly) and bear the same numbering scheme as on the Scud.
BTW, have you tried to measure the fin-span on the Safir? I get something like 2157-2184mm – corresponding VERY well to Scud-fins (span: 1810mm; body diameter: 880mm) on a NoDong-airframe (1250mm diameter; backed up by data in the UNMOVIC-compendium on the iraqi-north korean S-13-project, which in my opinion was identical with the Moksong/NoDong-A first observed in NK in 1990). Also pay attention to the number, size and arragement of the screws near the tip of the fins.
— Jochen Schischka · Feb 15, 03:49 PM ·
Hairs:
This discussion helps BIG TIME in understanding the evolution of the russian missile program.
Of course you are right – you can get even better protection by Nickel (i simply shunned that for being too expensive/“high-tech” for russian technology of the 50ies/60ies)!
After thinking about/researching on this a little bit, i’d say that chances are not bad that this IS exactly what was the difference between the Scud-B and C that made the reduction in wall-thickness on the 70ies- to 80ies-technology Scud-C possible!
The Scud-C obviously never entered series-production in the USSR, because a refit of the existing Scud-Bs would have been, as you correctly noted, expensive, and the much more capable solid-fueled 9M714/Oka/SS-23/Spider became available at about the same time (Scud-C probably even was some sort of “low-tech” back-up-programm for Spider -> similar maximum range, similar payload weight!).
As i wrote before, it would be VERY interesting to get a material-analysis on the Scud-C-sheet-metal from the Kuwolsan – just to leave not even a trace of doubt on that issue.
All in all, this could offer a plausible explanation HOW the Iranians achieved a dry-weight-reduction on the Safir (and perhaps also the Ghadr-1/NoDong-B)…but not necessarily without decrease in overall margin of safety. Let’s not forget that this missile is much longer than its predecessors (l/d of ~17.5 instead of ~12.5), so it would still be more “thin-skinned” in respect to flight at greater angles-of-attack or crosswinds due to the bigger lever of possible lateral forces (and there is also the possibility of a higher thrust-level)!
— Jochen Schischka · Feb 15, 05:12 PM ·
To Jochen Schischka:
Sorry I didn’t reply sooner; I just got back from a trip to Kennedy Space Center in Florida. I got to see a lot of the hardware from America’s first space vehicles. It was pretty neat.
> Your numbers do agree quite well with my own measurements/estimations,
> although i think that you overestimated the wall thickness by a factor
> of ~2, so the real number would be more in the order of 5-6mm
Thanks, I don’t have a good source for an estimate; I’ll use your figures. I’ve updated my estimate sheet:
Ratio Analysis
The main reason I want to estimate the expansion ratio is that I want to
estimate the chamber pressure, flow rates and turbopump particulars.
The new expansion ratio estimate is 9.7 +/- 0.5. Based on this expansion ratio
and the derived throat diameter, the optimal chamber pressure would be
about 58.3 Bar (845 PSI).
SAFIR STAGE 1 ENGINE MODEL
—————————————
Thrust: 291 kN (65500 lb-f)
Fuel: TM-185
Oxidizer: AK-27I
OF Ratio: 3:1
Pressure: 58.3 Bar (845 PSI)
Exp Ratio: 9.7:1
L* 175cm (68.88”)
Isp (SL) 243 s (ideal) 231 s (95% efficiency)
Isp (Vac) 270 s (ideal) 257 s (95% efficiency)
Isp (avg) 245 s (95% efficiency)
Unfortunately, these propellants don’t seem to fit the tank dimensions
I’m seeing (but neither do Kerosene/IRFNA). The best fit I’ve gotten is
with IRFNA/MMH but I’m still trying to fit dimensions. My current measured
tank dimensions are 5.58 m (18.3ft) and 7.16 m (23.5ft).
Do you know what technology is used for making Scud combustion chambers?
Are they brazed, electroformed or welded?
> There are also clear indications for an Isayev-heritage of this engine…
Yeah, it looks like the Iranian developers were very conservative and
chose to scale up an existing design. That’s smart engineering, I think.
Don’t innovate, just duplicate.
> BTW, have you tried to measure the fin-span on the Safir? I get something
> like 2157-2184mm
I’m not happy with the numbers I’ve gotten yet but I get 2138 mm (7.015 ft). It’s obvious
to me that they took a Scud tail and just attached it to the different-diameter
body.
Cheers,
— Ed LeBouthillier · Feb 21, 01:51 AM ·
> BTW, have you tried to measure the fin-span on the Safir?
> I get something like 2157-2184mm – corresponding VERY well to Scud-fins
I did a more accurate measurement on an image I captured from a video. Using the same error analysis that I used before, the fin span I measured was 2143 mm +/- 71.8 mm.
— Ed LeBouthillier · Feb 22, 01:27 AM ·
To Ed LeBouthillier:
Is it possible that you forgot to take the wall thickness into account for establishing the (inner) end diameter of the nozzle (627mm-2*5mm=617mm, +/- ~5mm)? Thus the Ae/At would be more like 9.6, corresponding to an optimum pc of ~55bar (according to my own estimations); Your values for Isp(eff) seem to be somewhat higher than my own (~248sec (vac) and ~224sec (SL)) – did you take the type of cooling (obviously film cooling with the fuel component), the open-cycle-bypass and the jet vanes into account? Is your thrust SL or vac (i get ~27t-SL, agreeing well with an exhaust-afterglow-length of ~16m photo-measured on the images of NoDong-A and -B released to the public so far)? For O/F i’d suggest the typical number for AK27I/TM-185 of 3.42 (like on the Scud), which brings me to the topic of the tank-volume-ratio:
Yes, this is somewhat tricky. Keep in mind that the Safir lower-stage is obviously a common-bulkhead design with an internal fuel line, so that alone distorts the visible ratio already slightly.
But there is another VERY interesting question to solve: Where is the pressure-gas stored? Since i estimate a requirement of about 300liters at 200bar, the normal amount of six 30l-pressure-bottles surrounding the engine (my working hypothesis for the NoDong-A, based on the Scud-B-configuration) would be insufficient. On the other hand, if the nozzles of the second-stage-engine really do have extensions (reasonable, since the nominal exit-pressure of the unmodified Serb-vernier-engine should be in the range of 1.5-2.0bar), then the space between the top of the upper tank and the staging seperation-plane is too small for a tank end-dome, the upper-stage nozzles AND a pressure-vessel. I think there is an additional, ~0.9m long compartment holding several toroidal Scud-C-pressure-gas-tanks between the engine-compartment and the lower tank (in my opinion, traces of this, aka a characteristical double-weld-seam, can be seen on some photos of the Safir-1), so the values i’m measuring with this in mind are ~4.2m and ~8.0m (i locate the intertank-bulkhead at about the same height as the handrail of the second working platform from the ground), matching IRFNA/Kerosene (as on Scud, NoDong-A and -B) quite good.
And since the NoDong-A can be, with some certainty, traced back to the soviet R-15-project, the NoDong-engine is probably not just a scaled-up R-17(=Scud-B)-engine, but more like the big ancestor of the 9D21 (indicating neither “innovation” nor “duplication”, but rather “license production”).
Some words on Scud (and with high probability also NoDong) engine technology: this is brazed steel-sheet-metal with an inlay of corrugated metal to form the cooling channels (like double wall corrugated board, just in 3D and steel…); the chamber is “baked together” in a vacuum furnace.
I hope this helps?
P.S.: Considering the fins: Did you try to compensate for perspective (another tricky business)?
— Jochen Schischka · Feb 22, 10:45 AM ·
To Jochen Schischka:
> Is it possible that you forgot to take the wall
> thickness into account for establishing the (inner)
> end diameter of the nozzle (627mm-2*5mm=617mm, +/- ~5mm)?
> Thus the Ae/At would be more like 9.6,
> corresponding to an optimum pc of ~55bar (according to my own estimations);
I tried to take that into account. But, the expansion ratio we both derive is about the same: 9.6 vs 9.7; they’re probably both within each other’s margin of error.
Here’s Nasa’s online engine simulator. I use my own math and software but this can be a useful tool for comparison (and other’s involvment):
Rocket Thrust Simulator
Here is my summary of what you’re suggesting:
Thrust: 27 metric tons-force
Expansion Ratio: 9.6
Nozzle Area: 2990 cm^2
Throat Area: 311 cm^2
Propellants: UDMH/IFRNA (used as a reference for the Nasa program)
Chamber Pressure: 5500 kiloPascal
Exit Pressure: 101.259 kiloPascal (sea-level pressure)
The derived thrust based on the above parameters is 27.1 kN.
However, if the launch vehicle is 26 tons then that would only give an acceleration of 1.04 G
(effectively 0.04 G). This is a very small take-off acceleration. My observations from film showed about 1.2 G acceleration (effectively 0.2 g). However, I so far have wide margin of error because the video quality isn’t too good. But, I measure about 4 seconds for the base of the vehicle to reach the level of the floodlights which are about the height of the first stage (52.5 feet or about 16 meters). Using basic physics math for distance and acceleration:
S = 1/2at^2
a = 2S/t^2
I derive:
a = 2*16/(4*4) = 32/16 = 2 meters/sec^2
Adding the effects of gravity, the effective acceleration is:
a = 2 + 9.8 = 11.8 m/sec^2 = 1.2 G
This means that the thrust has to be about 31.2 t-f (68800 lbs-f).
Using 1.04 as the acceleration, it would take about 8 seconds for the rocket to reach 16 meters of altitude. I think that there’s more thrust than 27 tons-force.
For the rocket motor to produce 31.2 tons with the established dimensions and its expansion ratio, the chamber pressure would have to be closer to 6425 kiloPascals (64.25 bar or 932 PSI).
> Your values for Isp(eff) seem to be somewhat higher than my own
> (~248sec (vac) and ~224sec (SL)) – did you take the
> type of cooling (obviously film cooling with the fuel
> component), the open-cycle-bypass and the jet vanes
> into account? Is your thrust SL or vac (i get ~27t-SL,
> agreeing well with an exhaust-afterglow-length of
> ~16m photo-measured on the images of NoDong-A
> and -B released to the public so far)?
We may be comparing apples and oranges. I’m looking at the combustion chamber as a component of a larger system. I’m only looking at the combustion chamber whereas you may be looking at the effective system performance (including all of the other source of propellant-losses like the turbopump).
Yes, those other factors would lower the effective Isp. I’m not considering those yet. My effective Isp is closer to the value you suggest.
I’m trying to work backwards using constraints based on observables to define the required engine parameters and then the system parameters.
> For O/F i’d suggest the typical number for AK27I/TM-185 of 3.42 (like on the Scud),
> which brings me to the topic of the tank-volume-ratio:
I’ll try those values. Thanks.
Regarding the tank volume ratio, I’ve been using the lift-strap indicators as likely partitions for the tanks. I show partitions at 2.33 meters (7.66 ft), 4.79m (15.70 ft), 6.96m (22.83 ft),
8.63m (28.83 ft), 9.49m (31.13 ft), 11.44m (37.52 ft) and 15.24m (50.00 ft) from the base of the first-stage cylinder (not the base of the tail fins).
I’m presuming that the actual tank junction will still be on these indicators (which may not be correct).
> Some words on Scud (and with high probability also NoDong) engine technology: this is brazed
> steel-sheet-metal with an inlay of corrugated metal to form the cooling channels (like double
> wall corrugated board, just in 3D and steel…); the chamber is “baked together” in a vacuum furnace.
Thanks. I haven’t been studying Scuds and missile technology so much as launch vehicles. My interest in the Safir has been as a launch vehicle.
Cheers,
— Ed LeBouthillier · Feb 22, 03:23 PM ·
To Ed LeBouthillier:
“the expansion ratio we both derive is about the same: 9.6 vs 9.7; they’re probably both within each other’s margin of error.”
Fully agreed.
“My observations from film showed about 1.2 G acceleration”
Well, i’m measuring ~1.3g, but as you correctly noted, the available material leaves some room for interpretation. And you’re absolutely right, if the liftoff-weight really was the officially proclaimed 26 tons (being roughly consistent with the upper end of typical densities for land-based missiles with storable propellants), then a thrust-SL of only 27t IS insufficient (i think we can agree on that this engine “looks like” it was initially designed for ~55-58bar and ~27-30t?), thus my idea with the “NoDong+”-hypothesis (of an uprated NoDong-engine with ~32-34t thrust-SL at ~65-69bar); Of course, another plausible explanation could also be that the takeoff-weight was only in the range of ~21-22t (lower end of typical densities). Or it could be a mix of both…i think we have to wait for additional evidence before we can satisfyingly solve this problem.
“Regarding the tank volume ratio, I’ve been using the lift-strap indicators as likely partitions for the tanks. I show partitions at 2.33 meters (7.66 ft), 4.79m (15.70 ft), 6.96m (22.83 ft),
8.63m (28.83 ft), 9.49m (31.13 ft), 11.44m (37.52 ft) and 15.24m (50.00 ft) from the base of the first-stage cylinder (not the base of the tail fins).”
If i subtract the ~0.4m from the base of the fins (i think this is the generally accepted reference for dimensioning missiles) to the lower end of the fuselage in my numbers, then i’m locating the tank-bulkheads at approximately ~3.1m (lower tank), ~7.7m (common bulkhead) and ~16.0m (upper tank) for comparison (with variations of up to -10%, depending on the particular photo); I used the double weld-seams (at least that’s where i think they are visible on some photos) as indicator for tank-end-domes/bulkheads (Don’t you think that SEVEN bulkheads are somewhat excessive aka weight-inefficient for only TWO propellants?). I’m observing blue-on white markers (labeled “SUPPORT” and/or “FIXING BAND” between two short, horizontal lines – indicating in my mind simply locations on the missile where the structure is strong enough for handling-purposes) at a height of ~2.2m, (~2.4m – you left that one out), ~5.3m, ~7.7m, ~9.8m, ~10.6m, ~12.6m and ~16.0m (i can’t find a marker at this last position), respectively. Again, more or less within each others margin of error, although my interpretation might be somewhat different.
BTW, these numbers were derived from pictures of Safir-1, since i haven’t found images of Safir-2 showing the I/II- or the III/IV-side of the missile (where the markers are painted on) yet.
Oh, and don’t get fooled by the mock-up displayed in Teheran in February ’08 and ’09 – this obviously non-functional dummy differs significantly from the flight-models (large NoDong-fins, illegitimately short cable-ducts, signs of an intertank section at an illogical place…).
— Jochen Schischka · Feb 24, 04:45 PM ·
> Well, i’m measuring ~1.3g, but as you correctly noted,
> the available material leaves some room for interpretation.
I probably should have said “At least 1.2 g…” Closer examination shows it to be >= 1.25g. But, we’re working with very sketchy information and having to deduce a lot from initial guesses.
> Or it could be a mix of both…i think we have to wait for additional
> evidence before we can satisfyingly solve this problem.
Agreed.
> i think we can agree on that this engine “looks like” it was
> initially designed for ~55-58bar and ~27-30t?
Agreed. If it was designed to operate at ~55-58bar, then yes it delivered ~27-30t. I’m not familiar with the technique you mentioned of measuring the flame length to derive pressure or thrust level. However, I would caution that this could be deceiving especially if they’re running fuel-rich or using film cooling to cool off the chamber walls. This always leaves a bunch of unburned propellant in the exhaust which burns externally.
If your NoDong+-hypothesis is that the engine is effectively ~32-34t at ~65-69 bar, then I think we’re both in agreement. My data supports an expansion ratio of 9.6 (or 9.7) with a chamber >= 64 bar. The major premises of this conclusion are:
1) the stated GLOW (Gross Lift Off Weight) is 26 metric tons
2) a presumed combustion gas gamma of ~1.22
3) a presumed combustion gas molecular weight of ~23.1
4) the observed combustion chamber expansion ratio (9.6)
5) the observed diameters of the chamber (Nozzle 617 mm, Throat 199 mm)
6) the observed acceleration (>=1.25g)
Of all of these, the “crux of the biscuit” is whether the stated GLOW is correct. Since we don’t have any better information, until we have evidence otherwise I think it’s prudent to accept that as true.
So, yes, I accept your NoDong+ hypothesis as I understand it.
> Don’t you think that SEVEN bulkheads are somewhat excessive aka
> weight-inefficient for only TWO propellants?
Yes. I don’t think all of them are bulkheads; some of them are likely strengthening ring locations.
Here’s a graphic of the model I’ve come up with so far (it’s still a hypothesis):
Safir Stage 1 Internal Layout
> BTW, these numbers were derived from pictures of Safir-1,
> since i haven’t found images of Safir-2 showing the I/II- or
> the III/IV-side of the missile (where the markers are painted on) yet.
There are several scans in the videos where they move the camera along the length of the vehicle where the markers are visible. I can make the stills available if you’d like; there’s about 2.5 Meg of images. I’m certain that they’re images of the Safir(2). I used VLC to capture video frames.
> Oh, and don’t get fooled by the mock-up displayed in Teheran in February ’08 and ’09
Definitely not. I wasted several days earlier before I realized that there were significant differences in design between the several versions. The actual Stage 2 that was launched appears significantly different than the Stage 2 that was shown in that hanger. The Safir(2) stage 1 also seems to have some differences from earlier vehicles. I’ve had to separate out all of my pictures by date to keep the analysis consistent.
Also, I have updated my Structural Coefficients Table (updated Scud B):
Structural Coefficients
Cheers…
— Ed LeBouthillier · Feb 25, 11:03 PM ·
“I’m not familiar with the technique you mentioned of measuring the flame length to derive pressure or thrust level. However, I would caution that this could be deceiving especially if they’re running fuel-rich or using film cooling to cool off the chamber walls.”
It’s a rough relation between the thrust (in kN) and the square of the length of the exhaust-afterglow (in meters); this works only in carbonaceous fuels, and it is rather resistent against respectively can be calibrated quite well to effects of film cooling etc. by comparison with similar engines with known data (in this case the Scud-B). With other known/estimated data on throat-diameter and cF you can deduce the chamber pressure. What actually IS a source of possible inaccurateness in this method is the not absolutely constant length of typical rocket flames (due to combustion fluctuations or mild POGO-oscillations), but otherwise it works astonishingly well – try out yourself (a +/-10%-fluctuation in thrust delivers a length variation of about +/-5%, and that should still be distinguishable).
Considering your “Safir Stage 1 Internal Layout”:
As i wrote before, there must be some place for additional pressure-gas-vessels on the first stage (don’t understimate the need for pressure-gas!). I suggest that there is an additional compartment for that between the engine-bay (which is obviously, apart from the fins, identical with the Shahab-3/Ghadr-1/NoDong – so there is no additional space for more pressure-vessels available there) and the lower fuel tank (shortening that one for ~0.9m).
“I can make the stills available if you’d like; there’s about 2.5 Meg of images.”
I’d be VERY interested in these – although i’m always somewhat skeptical in respect to data derived from video-stillframes (from experience, i’d say that it’s even harder to accurately compensate for perspective in such images than in standard photos).
Considering the Safir mock-up:
With my previous comment i was rather hinting at the outdoor-display with the erected missile on the MEL; I tend to the interpretation that the August-08-display inside of that hangar showed the REAL Safir-1, but in disassembled/delivery state (for example, the retro- or ullage-rockets weren’t mounted yet, the jet vanes were still missing, several red-coloured protective coverings were still in place, etc.).
BTW, i think i found some errors in your Structural Coefficients Table: The Aggregat-4/V-2 was NOT a pressure- but a pump-fed design; and i think you combined the empty weight of a late version with the propellant weight of early versions (late versions had a propellant weight of 9025kg/19897lb while early versions had an empty-weight of 3008kg/6632lb or more);
In contrast, the Scud-A/R-11 was pressure- not pump-fed (it was obvously derived from the quite similar german Wasserfall-SAM);
And is it possible that you forgot to incorporate the 15kg/33lb of pressure-gas into the empty-weight/fueled-weight of the Scud-B/R-17? According to my sources, this missile also was fueled with only 3771kg/8314lb.
— Jochen Schischka · Feb 27, 10:35 AM ·
To Jochen Schischka:
> As i wrote before, there must be some place for additional
> pressure-gas-vessels on the first stage (don’t understimate
> the need for pressure-gas!).
Yeah, I haven’t gotten that far yet. However, some of it might be generatable from propellants (I’m not saying that I believe that’s what they’re doing; I’m just willing to entertain the
possibility right now).
> I suggest that there is an additional compartment for that
> between the engine-bay (which is obviously, apart from the
> fins, identical with the Shahab-3/Ghadr-1/NoDong – so there
> is no additional space for more pressure-vessels available
> there) and the lower fuel tank (shortening that one for ~0.9m).
Yeah, I haven’t worked on that yet. I’m still trying to fit the propellant volume into the tanks properly. It’s been harder than I imagined.
As an example, here are where I see the partitions in tabular form;
I’ve labeled them T01, T02, T03, T04, T05:
Tank Partitions
Based on the above table, I searched among likely propellants for reasonable combinations that would meet several properties:
1) The propellant density was greater than 83 lbs/cuft (1.33 g/cc).
2) The propellant impulse density was greater than 20000 lb-f-sec/cuft
3) The propellant combustion temperature was less than 5000 deg F (2760 Celsius)
4) The propellant volume ratio matched those of the above partitions
I’ll explain these requirements if asked.
First, I needed to come up with some likely propellants and their properties. Here’s the list after applying the requirements:
Selected Propellants
The results were surprising to me. I didn’t expect the results that I got. Based on the above search, the most likely propellants are:
AK-27I with Tonka 259 at a ratio of 3.2
IRFNA with Aniline at a ratio of 2.7
I select AK27I with Tonka as the most likely (but I could be wrong). Do you have an opinion on this?
Here’s an ISP and temperature table versus the oxidizer to fuel mixture ratio. This was derived using Propep:
ISP and Temperature
Basically, the model that I showed the other day still holds (which also surprises me). The likely physical layout is still:
Possible Physical Layout
One reason that I don’t like this outcome is because it presents some problems: the fuel pump must be connected to the lower tank. The fuel pump is likely delivering higher pressure because of the higher density of the propellant. It also likely increases the amount of horsepower (kilowatts) to pump the propellant. But these are not discounting problems.
> I’d be VERY interested in these
Here is a link to a zip of the files:
Captured Images
I’ll keep these available at the above link for one week.
Anyone else wanting them in the future can contact me.
> BTW, i think i found some errors in your Structural
> Coefficients Table: The Aggregat-4/V-2 was NOT a
> pressure- but a pump-fed design;
You’re right. It’s a silly error on my part. I decided that I will maintain this document with versioning because I find it useful. I welcome any comments or questions related to that table. Thanks for the corrections.
> And is it possible that you forgot to incorporate the 15kg/33lb
> of pressure-gas into the empty-weight/fueled-weight of the Scud-B/R-17?
I had to make a decision whether to consider it part of the empty weight or the consumable propellants. I generally tried to include it as consumable propellant (if I recall properly). But, I also have poor sources of information on the Scud B so these are just my best estimates. I’m trying to maintain references on these figures and I’ll consider suggestions. I think a table like that is useful to myself and others, so I’m going to maintain it for some period of time (several years into the future).
I’ve uploaded a corrected (and dated) version of the file at the same link as before.
Cheers…
P.S. I apologize for the non-metric values in some of the above stuff, but I didn’t have the time to convert everything.
— Ed LeBouthillier · Feb 28, 09:13 PM ·
Ed:
The first entry in your structural table, Atlas E with booster, can’t possibly be right. It could be right for Atlas E without boosters, only sustainer.
— Murray Anderson · Mar 1, 03:03 PM ·
To Ed LeBouthillier:
First of all, thank you very much for making your captured images (and the other photos) available to me.
BTW, the pictures “66a.jpg”, “247.jpg”, “9777.jpg” and “SAFIR_26a.jpg” show exactly the non-functional mock-up i was commenting on before, and “Omid-pre-launch.jpg” shows Safir-1. Have you noticed that the Iranians obviously rotated the missile 90° on the launching-table (compare for example the numbering on the fins in pictures “vlcsnap-12193606.png” and “vlcsnap-12191771.png” – the latter also clearly shows the jet vanes)?
“However, some of it might be generatable from propellants” (pressure gas):
Hmmm, i can’t exclude this possibility, although the level of sophistication indicated by that would be, to put it mildly, really amazing (and maybe this would also require additional space for gas-generators…at least additional valves and fuel lines). Since i’m a big fan of Occam’s razor, i’d prefer an explanation as simple as possible utilizing components verifiably available to Iran (and North Korea?) like toroidal pressure-gas tanks from the Shahab-2/Scud-C (~60l, 200bar, ~0.2m high, ~0.8m diameter).
Considering the tank partitions:
As i wrote before, in my opinion there is an additional partition between “H” and “G”; I think in captions “vlcsnap-12195435.png”, “vlcsnap-144804.png” and the photo “X00898108700.jpg” traces of a characteristical double-weld-seam can be perceived, although i must admit that this is really hard to see; I suspect that the Iranians spackled the corresponding area up – compare for example the area around “F”, where i think the common bulkhead between both tanks can be located with some certainty (if this is not the case, what is otherwise the purpose of the service-platform slightly below that region? Also, it could be seen more clearly on pictures of the Safir-1), in pictures “vlcsnap-176451.png” and “vlcsnap-176577.png”. Otherwise, i’d choose configuration T04 as the most likely. (A little suggestion on the side: maybe your tank-end-domes are slightly too shallow; based on what i know of the Scud, i’d expect a height of ~0.3m.)
Considering your suggestion of an IRFNA/Tonka-combination (well, AK-27I is some form of IRFNA, and Tonka is also a mixture of Aniline and Xylidine…):
The almost identical layout of the engines of NoDong and Scud-B, the almost identical type of exhaust-plume and the visible tank-volume-ratio of the NoDong-A lets me come to the conclusion that the standard NoDong-engine uses the same propellants as Scud-B: AK-27I + TM-185.
Do we agree that the engine compartment of Safir IRILV and Shahab-3/NoDong-A or Ghadr-1/NoDong-B do look VERY similar (and thus with high probability also contain more or less the same engine)?
A redesign of an existing engine for a different propellant combination would require at least a COMPLETELY new turbopump-array, different regulation valves etc. (due to other densities, different optimum-O/F etc.) and several extensive changes to the thrust chamber (e.g. the number/arrangement of injection nozzles); Due to different thermal properties of the new fuels, a radical change in the cooling concept could also be necessary – all in all VERY work-intensive and challenging.
On the other hand, a simple uprating of the chamber-pressure (adequate design-inherent reserves provided) would be rather straightforward: As far as i see, this could be accomplished by either a slightly higher adjustment of the regulator valve for the tank-pressure or by a higher rotation speed of the turbopump due to a modification to the thrust regulator valve (aka larger mass-flow into the gas-generator).
That said (or rather written), Isayev also designed engines with the propellant-combination AK-20K/TG-02 (for example for S-75/SA-2/Guideline, S-200/SA-5/Gammon, R-13/SS-N-4/Sark or R-21/SS-N-5), so i’d try this if you want to follow this hypothesis further (i think there can be no doubt of the Isayev-heritage anymore).
BTW, if you like, i can provide more information on R-17/Scud-B and Aggregat-4/V-2 (but the sources on both are in german).
— Jochen Schischka · Mar 1, 05:14 PM ·
To Jochen Schischka:
> BTW, the pictures … show exactly the non-functional mock-up i was commenting on
> before, and …
Ah, that makes sense. Thanks.
[Regarding fuel generated ullage pressurization gas…]
> Hmmm, i can’t exclude this possibility, although the level of sophistication
> indicated by that would be, to put it mildly, really amazing
Yeah, you’re probably right. I haven’t begun to think about pressurization yet, though.
> As i wrote before, in my opinion there is an additional partition between “H” and “G”;
I guess it’s possible. I’ve looked and haven’t seen anything. Also, if
you’re right about AK-27I + TM-185 propellants, then it won’t be necessary
to have a partition there. It could just be a holder for the oxidizer feed
through the tank.
With AK-27 and TM-185, the propellant tanks would be about 8.26 m (27.2 ft)
and 4.60 m (15.1 ft) and would likely extend beyond any partition
between H and G.
> The almost identical layout of the engines of NoDong
> and Scud-B, the almost identical type of exhaust-plume
> and the visible tank-volume-ratio of the NoDong-A lets
> me come to the conclusion that the standard NoDong-engine
> uses the same propellants as Scud-B: AK-27I + TM-185.
It’s definitely one of the likely ones. The one reason that
I was steering away from it was because I was having trouble
fitting the propellants plus ullage space into the tank lengths
that I calculated. That’s why I like the AK-27I + Tonka 259.
I have trouble fitting AK-27I + TM-185 into the tanks and
leaving some ullage space. But, maybe I’m wrong about the
distance between the markers.
Maybe I have to lengthen my first stage estimate a bit. I’ll
try to do a better analysis of both the overall length (with
error margins) and the distance between the partition markers
(with error margins). But, the history of AK-27I + TM-185
with the Scud probably makes it a more likely propellant.
Also, in reviewing my table, I found a few errors.
First, was a minor density error which didn’t effect anything
and, second, was a major review of my error analysis. I think
I was too lazy in assessing the tolerance for the volume ratio
and the results suggest several more likely propellants:
IRFNA + Kero with a mixture ratio near 3.00
AK-27I + TONKA-259 with a mixture ratio between 2.80 to 3.20
AK-27I + TM-185 with a mixture ratio likely between 3.5 to 4.0
AK-20 + TONKA-259 with a mixture ratio near 3.20
I did a slightly better error analysis of the tank lengths
(propagating errors from my previously calculated lengths)
and used that to better select possible propellants. I
presumed 5% error in length measurements (which is consistent
with what I’ve been able to get before) but I don’t have any
error estimates for propellant densities or combustion
characteristics. Unfortunately, it doesn’t pin anything down
any better because there are several different options still.
Here’s the updated propellant and tank study:
Propellant Study
> BTW, if you like, i can provide more information on R-17/Scud-B and
> Aggregat-4/V-2 (but the sources on both are in german).
That would be greatly appreciated, thanks. I’m fine with German.
My email is: codemonky .at. earthlink.net
I’m going to start considering pressurization. I think that the
propellant is narrowed down to a few possible fairly dense
propellants with very similar properties.
Cheers…
— Ed LeBouthillier · Mar 2, 08:19 PM ·
To Murray Anderson:
> The first entry in your structural table,
> Atlas E with booster, can’t possibly be right.
> It could be right for Atlas E without boosters,
> only sustainer.
First, thanks for your commentary.
I don’t have the precise figures that I used to calculate those
values. When I first started the table, it was for my own use and
I was sloppy in maintaining records. But, I’ve tried to use the
best sources I had available (and it is possible that I’ve
misunderstood something).
But, using Nasa Document 19780071276, Launch vehicle handbook, Aug 11, 1961,
for the Atlas E ALWASS, I come up with figures very close to those. The
document is accessible here:
http://ntrs.nasa.gov/search.jsp
They list the following values (I’ve added the propellant weights together):
Initial Weight: 262,564 lbs
Prop. Weight: 248,640 lbs
SBW 7232 lbs
SBW + Jettisonable: 7455 lbs
(SBW = stage burnout weight)
These are within a small percentage of the
values that I published. In my table I used
the value of 250,361 lbs for the propellant
but this shows 248,640. Even using the
lower value for propellant, the Structural
Coefficient is still 0.029.
So, as far as I can tell, it’s correct.
Now, though, I have to admit a condundrum. There
may be an error in my figures because the same
document also reports that the Stage Propellant
Fraction is 0.948. This figure would be the
conjugate of my figure, and would suggest a
number more likey 0.052. So, I may have to
re-analyze things.
What do you think the correct answer is and
how should I derive it?
I welcome any commentary, though, and I will
endeavor to better document the numbers I’ve
used. I’m almost thinking that I may have to
make a booklet with supporting figures and
source references for this to be of the kind
of value that I think it should be.
Cheers…
— Ed LeBouthillier · Mar 2, 09:10 PM ·
Ed:
The source I’m using is http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19780071276_1978071276.pdf , page H3. It gives (7455+6469)/262564 = 0.053 as the structural mass fraction. The 7455 lbs is the two booster engines and associated stuff, discarded at 120 seconds, and the 6469 lbs is the sustainer engine and tank weight.
The table is extremely difficult to read and comprehend, unless you keep the odd structure of the old Atlas in mind.
— Murray Anderson · Mar 3, 12:01 PM ·
To Murray Anderson
> The source I’m using is … and the 6469 lbs is the sustainer engine and tank weight.
Thanks for pointing out the error. I’ve made what I believe are the proper corrections and also provided supporting data. I’ve decided to keep a webpage dedicated to this table:
http://home.earthlink.net/~apendragn/atg/coef/
I think I got it right this time. I’m also going to start reviewing other less-certain rocket weights.
Cheers…
— Ed LeBouthillier · Mar 3, 11:44 PM ·
Although this is not a very different photo from many others, this Washington Times photo of the Safir(2) is fairly large:
http://media.washingtontimes.com/media/img/photos/2009/02/03/20090203-145341-pic-701023937.jpg
You can see weld lines and all sorts of detail in this 2334 × 3500 image.
— Ed LeBouthillier · Mar 5, 11:24 PM ·
To Ed LeBouthillier:
Excellent picture – thanks!
I still think there is a certain amount of uncertainty considering the lower end of the lower tank of the first stage; note for example the weld lines on the upper stage and how almost invisible (concealed?) the double weld-seams of the tank partitions are in spite of the high resolution. (I would caution against using ONLY the painted markers on the missile as indicators: on the Scud-B, only one out of five markers on the cylindrical section of the body actually coincides with a tank-bulkhead – and i haven’t seen striking evidence for a bulkhead directly above the engine on the Safir yet…)
By now i think that perhaps additional 10-(Scud-B) or 30-liter-(NoDong)-airbottles were used, since a toroidal 60l-tank (Scud-C) in a compartment between engine-bay and lower tank would possibly obstruct the propellant lines, although that would be a far less weight-efficient method; Perhaps there is a ~150l-torus available from some other old soviet missile system?
Yet another possibility could also be pressure-vessels INSIDE of the fuel tank (on the upper end, around the oxidiser-line), but that would with high probability prove to be a maintenance-nightmare in case of some sort of malfunction.
BTW, did you get my email?
— Jochen Schischka · Mar 9, 01:33 PM ·
To Jochen Schischka:
> I still think there is a certain amount of uncertainty considering the lower end of
> the lower tank of the first stage;
I agree. I don’t have any good pictures of that part. My error distribution is quite
large in the lower area. I guess we’ll have to wait for another satellite launch
and hopefully get better pictures.
> note for example the weld lines on the upper stage and how almost invisible
> (concealed?) the double weld-seams of the tank partitions are in spite of the
> high resolution.
What does the double weld signify? What process creates it?
> (I would caution against using ONLY the painted markers on the missile as
> indicators: on the Scud-B, only one out of five markers on the cylindrical
> section of the body actually coincides with a tank-bulkhead – and
> i haven’t seen striking evidence for a bulkhead directly above the engine
> on the Safir yet…)
Yeah, I’m starting to question some of my own presumptions.
Up until that last picture, I didn’t have anything clear enough to suggest
anything else, but now I do have to doubt my earlier presumptions.
I’m also questioning some of my earlier rocket parameters because simulations
have shown fairly high gravity losses because of the relatively low takeoff
accleration. I see the delta V requirements looking like this:
Orbital Velocity: 7.76 km/s (25456 ft/sec)
Gravity Loss: 1.43 km/s (4700 ft/sec)
Aerodynamic Loss: 0.08 km/s (270 ft/sec)
Extra Margin: 0.15 km/s (500 ft/sec)
———————————————————————-
Total DeltaV: 9.45 km/s (31000 ft/sec)
Realistically, the vehicle probably needs a little bit more margin than
I’ve allowed above. Currently, I’m presuming 9.6 km/s (31500 ft/sec).
I don’t want to add unrealistic delta-V; I’m trying to creep up on a
lower bound for performance.
So this implies a somewhat higher performing vehicle than I was originally
thinking about. This necessitates lowering the structural coefficient
below 13% for the second stage and below 14% for the first stage. This is
still in the typical pressure-fed vehicle range (not high tech) but better
than I originally thought it would be. It also indicates significant improvements
over basic Scud engineering.
> By now i think that perhaps additional 10-(Scud-B) or 30-liter-(NoDong)-airbottles
> were used, since a toroidal 60l-tank (Scud-C) in a compartment between engine-bay
> and lower tank would possibly obstruct the propellant lines, although that would
> be a far less weight-efficient method; Perhaps there is a ~150l-torus available
> from some other old soviet missile system?
I agree with you here too. I did a model where 16 30-liter bottles could fit fairly
nicely into the base. There’s a lot of room around the engines and this IS the
typical Scud approach:
Scud B Tanks
> Yet another possibility could also be pressure-vessels INSIDE of the fuel tank
> (on the upper end, around the oxidiser-line), but that would with high probability
> prove to be a maintenance-nightmare in case of some sort of malfunction.
I think that the simplest solution is just Scud tanks; there’s plenty of room for them,
it seems to me. They could be arranged in several different ways, including clusters
to accomodate the space available. There’s a big problem with gas expansion and
cooling that can really only be solved by having a heat source (i.e. the engine)
close by. If they place the tanks anywhere else, they need to transport heat to
wherever the tanks are (or have a heat source there).
> BTW, did you get my email?
I did and I’ve responded several times but email keeps bouncing back. Do you have
a more reliable alternate email address? I don’t usually have problems to Europe.
In one of the emails I was suggesting that the end caps were likely torispherical:
Torispherical Endcaps
Torispherical
Here are images of Scud end caps which confirm this design:
Scud Endcaps
— Ed LeBouthillier · Mar 12, 11:07 PM ·