F-14AM - The Iranian Tomcat - History, Performance, and Discussion

Let me just sum up what you’ve done on this topic so far.
“My math for figuring out thrust where I don’t take drag into account and also have no clue what the mass is of the object my sustainer is powering is actually always correct and everyone else is always wrong and biased. Also @sudo_su1 your sources aren’t average thrust. You’re wrong. simple as.”
NASA: “They’re average thrust.”
“Well then they were launched at altitudes in this particular case.”
NASA: “They’re ground launched.”
“Well then there were different atmospheric conditions. Also NASA did their calculations wrong. My single sample size and napkin algebra is better than all of NASA’s launches and their team’s calculus combined.”

You’ve done nothing but endlessly argue completely ridiculous points from a baseless position. For the benefit of everyone just never comment again.

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Sudo pushed a report before all the facts were at hand and before all the questions could be asked. He has since revised his post dozens of times with new information and for some odd reason, only the information that agrees the missile should be buffed. This is a cue for speculation imo. Dismissing useful information when it doesn’t agree with you is NOT proper reporting etiquette.

The ROTEX-T launch mass is known, in fact they broke down the components for us as well. The drag can be found by Gaijin rather easily as the rocket uses a very simple shape and the deceleration post-burn of the initial booster stage (PRIOR to M112 ignition) gives us a solid base from which to ascertain the true drag coefficient of the rocket.

Not only that, none of the NASA sources have a true thrust over time chart to go off of. Gaijin can use the various average thrusts from the NASA source in conjunction with the thrust profile provided by the ROTEX launch to find the true minimum and peak thrusts as well as determine how they want to translate that into the game.

You want to paint my bickering as baseless but instead what you’ve done is come in here with an unwarranted spiteful insult and added precisely nothing to the thread or conversation. If anyone was to see themselves out, it’s you… but this is a public forum so we’ll all just have to suffer with each other won’t we?

@sudo_su1 pending NASA’s reply about how they found the “average” thrust… I’m sure they put a lot more effort into it than this but as an example…

How many times did they sample the thrust value? At what intervals? Did they take absolute peak and absolute minimum to divide by 2? Without knowing how they came up with the data our best bet is to assume it is well-sampled and apply the average to the thrust curve shown in the ROTEX-T source.

Specific impulse and burn rate isn’t constant, so in order to calculate thrust from acceleration you need to also estimate / calculate vehicle’s mass (which includes the mass of the remaining propellant) at each point in time.
Good luck with that with the amount of info that the ROTEX-T study provides.

And actually, the very fact that the authors of the ROTEX-T study themselves didn’t even try to calculate and provide a thrust figure, shows that they understood that they don’t have enough data to do so.

There’s no need.
All you need is booster burn time * average thrust and sustainer burn time * average thrust
That’s all that WT models anyways.

Plus, how many of the rocket motors that we have in the game have actual thrust charts?! :)

The idea that a single launch in Europe where the authors themselves didn’t bother with calculating the thrust figures most likely because they understood that they don’t have enough data to do so, is a better source than NASA’s reference handbook that has been refined and revised over many many years using the data from many many launches, is absurdly childish.

We know the weight of the initial booster section, the payload, the weight of the M112 rocket motor section. Subtract the initial booster and then it’s as simple as determining boost - sustainer mass fraction.

Adjusting the weight at booster burn-end using the average thrust values in this chart until we have a reasonable ISP for both booster and sustainer gives us reasonably correct values and propellant burn rates.

From there it’s easy to determine how much mass is burned over time. The data is there, why are you so adamant to ignore it?

R-27ER, and even then it’s underperforming in top speed by 100 m/s at altitudes 5km+

Because that wasn’t the objective?

Certainly NASA’s handbook has an acceleration over time chart that shows us the exact thrust profile of the rocket motor? No? Then yes this source is 100% useful. Stop pretending it’s one or the other, it is both. All of these sources together. Not just one. Quit dismissing useful data.

Your best friend ranting about my math and providing nothing useful to the convo was childish. Let’s not turn this into a slapping contest.

Devs responded about Fakour thrust
https://community.gaijin.net/p/warthunder/i/EaLuXWmNnW4F?comment=KhmvL8Y3rLGxArCHRfuLqUqh

What I didn’t realize - that the devs caught - is that the specific impulse of ~234s is given for the rocket type in one of the sources and explains that a specific singular type of propellant was assumed to be used for the approximation of the thrust values… leading to incorrect figures.

Simply put, it was not (and still isn’t possible) for a rocket with AP/Poly fuel to have nearly 300s specific impulse. The numbers in-game are much closer to the real values than the NASA sources (which approximated average PEAK thrust) are giving.

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Yeah they responded without even properly reading the report!

Completely false argument due to the fact that they didn’t even read the article … They just searched “specific impulse” in the article and copy pasted the first number they found …

Refer to the section “3” in my next post.

But the TLDR of it is that this study literally takes the thrust figures from the 2005 edition of the NASA handbook and calculates the specific impulse off of that.

So of course the specific impulse will be in line with the thrust figures of the 2005 edition of the handbook! what a shocker!

Regarding this bug report: Community Bug Reporting System

Which was answered as “not a bug” by the developers, unfortunately, as is clear by the dev’s response they did not look into the evidence thoroughly:

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Таким образом, очевидно в этом источнике тут указаны пики тяги, средняя тяга за время горения РДТТ M112 значительно ниже.
Как выше я уже сказал это не корректные данные, вы аппроксимируете пиковую тягу на все время работы разгонной и маршевой ступени. Для РДТТ достижение суммарного импульса в 855kN при массе топлива в ~295кг не возможно, потребовался бы удельный импульс 2900м/с, чего у РДТТ не достижимо в условиях земной атмосферы.
Где в источнике сказано что это средняя тяга?
Это не более чем ваше предположение. Которое опровергается значением удельного импульса двигателя.

I have already emailed NASA and they confirmed that the values in the handbook are in fact average thrust, and not peak thrust:

Spoiler

And this is obvious anyways … Peak thrust is useless to the audience of the NASA rockets user handbook who are looking to choose an appropriate motor for their application. The source explicitly mentions “average thrust” for all other motors that are listed in the handbook. For the “Improved Orion” motor it uses the term “approximate thrust” because it doesn’t provide “exact” values i.e. to the first digit, like it does for other motors.

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Кроме того это противоречит источнику, который вы сами приложили к репорту:
Imgur: The magic of the Internet

The study in question is this:
REXUS 2 - the first EuroLaunch project
https://www.researchgate.net/publication/238585597_REXUS_2_-_the_first_EuroLaunch_project

This study is from 2005. And:
A- Does not provide any thrust figures.
B- There is no indication of what measurements were taken during flight. It does not even present any acceleration charts. The only thing it provides is, not even an actual flight trajectory chart, but a “predicted” flight trajectory chart.
C- The study explicitly specifies that the burn time given is not the actual burn time, but the “nominal” burn time:

Spoiler

https://i.imgur.com/31et2DL.png

So it’s obvious that the burn time mentioned in the study is a quote from other older sources (and is also close to what is mentioned in the 2005 edition of the NASA rockets user handbook). Not to mention that NASA’s 2023 rockets user handbook is a more recent source / edition and has revised information.

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Если считать что эти данные корректны то суммарный импульс тяги то получается (200006+4000 18) * 0.4536 * 9.81 = ~854364Н * с, что не возможно для РДТТ с массой топлива в ~290кг. В источнике который вы приложили указано что удельный импульс для этого РДТТ составляет 2305м/с
Imgur: The magic of the Internet

As for the specific impulse, the study that the devs are quoting here (Multidisciplinary optimisation of single-stage sounding rockets using solid propulsion) is not even a test launch study, but rather a simulation study.

And this study is not even simulating the propellant used in M112, but rather an alternative propellant based on other propellants that they had simulation models for (from other, different rocket motors):

a slightly modified propellant, closer in performance to the existing ones, was used
(specific impulse of 2305 m/s at launch at a maximum chamber pressure of 14 MPa, one type of propellant in the
SRM is assumed).

Moreover, this study itself, takes the thrust figures from the 2005 edition of the NASA rockets user handbook!:

Spoiler

https://i.imgur.com/LMkIR8e.png

Reference [33] highlighted in the screenshot above is (The Improved Orion section of) the 2005 edition of "NASA rockets user handbook".

And the source itself says:

(specific impulse of 2305 m/s at launch at a maximum chamber pressure of 14 MPa, one type of propellant in the
SRM is assumed

“is assumed”! is the key word here!

I.e. this study literally references and takes the thrust figures from the 2005 edition of the NASA handbook and calculates / assumes the specific impulse off of those values.
So of course the specific impulse calculated / assumed based on the thrust figures give in the 2005 edition of the NASA handbook will be in line with the thrust figures of the 2005 edition of the handbook! And lower than that of the 2023 edition of the same NASA handbook (which specifies higher thrust figures). What a shocker!

Moreover:

A- According to the 1960 patent by the manufacturer (Aerojet) which seems to be the first patent regarding AP/polyurethane propellants, it’s said that depending on the aluminum content, the specific impulse can range from 220 to 270 s at 1000 psia:

Solid propellant with polyurethane binder US3793099A
https://patents.google.com/patent/US3793099A/

Spoiler

https://i.imgur.com/0uUgjx7.png

B- Later, a 1963 patent by the manufacturer (Aerojet) increases this figure to about 300 s (for comparison, values given in the 2023 edition of the NASA rockets user handbook will result in an Isp of around 295.6 s):

Polyurethane propellant formulations and process US3291660A
https://patents.google.com/patent/US3291660A/

Spoiler

https://i.imgur.com/gi7tKLm.png

4-
Even if we assume that the actual specific impulse of the propellant can’t be that high, we already have motors such as that of AIM-7F that have in-game specific impulse values higher than “physically possible”.
The reason that the devs gave for this was that:

https://community.gaijin.net/issues/p/warthunder/i/RfdZe2n1F4OA
The over performing motor is intentional, when a missile’s motor is firing it will experience a reduction in drag compared to when motor is not firing. Since it is difficult to implement two different drag value, the dev decides to increase the thrust during the motor firing to simulate the drag reduction.

So it’s possible that the thrust values calculated in the 2023 edition of the NASA’s rockets user handbook are also high as a result of the drag reduction that happens due to the long burn time of the motor.
And if that’s the case, there is no reason as to why that shouldn’t be reflected in the game as well (just as it’s reflected for the AIM-7F by giving it specific impulses higher than “physically possible”).

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В этом источнике очевидно речь идет не о М112, т.к. для достижения импульса в 934720Н*с при массе топлива в 295кг потребовался бы удельный импульс 3168м/с, чего физически не возможно для РДТТ. Кроме того там указана одна тяга, это означает что используемый во второй ступени РДТТ однорежимный либо данные указаны не полные или вообще ошибочные.

The study in question is this:
A sounding rocket payload experiment on zero gravity fuel gauging using modal analysis
https://www.researchgate.net/publication/273308323_A_sounding_rocket_payload_experiment_on_zero_gravity_fuel_gauging_using_modal_analysis

Improved Orion is an off the shelf military surplus M112 motor used for space / sounding rocket applications as indicated by many studies.

There is no “sustainer-only Improved Orion / M112” motor.

The reason that only one thrust value and one burn time is given is because the researchers averaged the total impulse over the total burn time to calculate one overall average thrust over the whole burn time, without specifying each phase separately.

The reason that the specific impulse is so high is because here the M112 is burning as second stage, at very high altitude, which results in higher specific impulses.

But this confirms the burn time and average thrust figures that were revised in the 2023 edition of the NASA rockets user handbook, because, if the burn time for a second stage very high altitude burn is 25.4, it can’t be 26s for a first stage burn from the ground.

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If you wrote smth in comments instead of report itself it can be easilly missed, especially if there is a long conversation between different players and not just tech mod and you.

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Well, I sent all the important comments including the email from NASA confirming that the figures given in the handbook are average, rather than peak thrust, to the tech mods (in DMs) and asked that they be included and looked upon.

So I’m assuming either those weren’t passed or the devs didn’t bother to read them. Hence the dev’s response that “It’s just your assumption that those values are average thrust”.

Though I’m equally shocked by the fact that the devs think just because that one study gave the average thrust for the total burn time without specifying each phase separately, it means that the study is wrong! Or that there is some other single-stage sustainer-only Improved Orion / M112 motor!

You and I both missed important information lmao

Looks like my screenshot highlighting some portions of the document that I sent them.

This is erroneous or calculated improperly.

This is true for every rocket ever made.

Which are more credible / accurate and confirmed by ROTEX-T report accel chart.

It’s valid, you didn’t even read it.

You’re ignoring the part where the 2023 NASA handbook ISP is not physically possible irl with that propellant. I told you this already.

Never seen in any production vehicle. Just a possibility not realized due to other factors such as storage longevity and reliability requirements.

Paper often doesn’t translate 100% to reality.

Citing a missile in-game with propellant masses switched around erroneously according to primary data isn’t going to validate your nonsensical conclusion that AP/Poly must outperform every propellant type made after it until the mid 2010’s.

??? Doubtful. I think they just lied or didn’t calculate it properly but guesswork as to how is not useful.

Which is erroneous and even if correct isn’t particularly useful to find thrust curve.

They read them, I also forwarded a lengthy reply explaining your bias and forwarding ALL of the available evidence.

As I have said like a 100 times already, the ROTEX-T study uses the M112 as second stage, for a very high altitude burn.
So the burn time is expected to be longer.

Yeah, NASA is lying or stupid.
They should hire you for their thrust calculations!

A study that is not even presenting any acceleration charts, nor even any actual flight trajectory charts (which is the simplest of things to present) and instead presents a “predicted” flight trajectory chart.
And doesn’t even claim that the 26s is the actual burn time, but rather uses the term “nominal burn time” which essentially means that this is what is usually stated for this motor, cannot be used to invalidate NASA’s actual test data.

Yeah, people from the University of Oregon, Carthage college and NASA didn’t know how to calculate the thrust, but you do :)

There’s absolutely no need for a thrust curve.
The game doesn’t even model a thrust curve. It only models burn time and average thrust.
And there is no public thrust curve for most rockets in the game anyways.

What’s valid?
It’s literally using the thrust values given by the 2005 edition of the NASA handbook. (reference [33] in the article)
So the 2023 edition of the handbook is invalid because some study has used the thrust values of the 2005 edition to calculate the specific impulse?! :)

The source itself says “specific impulse of 2305 m/s … is assumed”! That’s the key word … “assumed”!

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You’re so blindly batting for a buff and ignoring evidence to the contrary. Don’t be surprised if it doesn’t get the results you want. You can’t ignore the fact that 295s impulse is literally impossible.

Yeah, “evidence to the contrary” :)

An article which is essentially saying: “Based on the thrust figures given in the 2005 edition of the handbook we assume that the specific impulse is …”

This is evidence that the revised thrust figures in the 2023 edition are wrong! :)

Well, according to you NASA is 100% infallible and couldn’t possibly accidentally claim a 1960s ammonium perchlorate propellant could achieve 295s impulse ASL…

You are yet to properly address how the I-HAWK motor could have such an implausibly high ISP.

And no “the AIM-7F is implemented incorrectly so the HAWK should be too” is not a valid answer.

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I’m not saying AIM-7F is implemented incorrectly.
I’m saying if AIM-7F has higher “apparent” thrust due to the drag reduction that the long burn of the motor causes, there is no reason why the same shouldn’t be reflected for the M112 as well.

That could be one of the possible explanations of why the thrust figures given for the M112 are relatively high.

And who says it’s implausible?
What’s your source?

There are two sources that confirm such high Isp (NASA’s handbook and the study in the section C of the main report)
And manufacturer’s patent also claims such high Isp.

So what’s you source on it being “implausible”? Your gut feeling?

Even the AIM-7F (1976), AIM-120 (1991), and other newer missiles do not exceed ~260s in surface conditions. It’s not remotely feasible - patents exaggerated claim or not.

It is VERY well known that the efficiency of solid rocket propellants did not exceed 270s ASL until the 2000s in actual applications.

Those motors use not only different propellants but different types of propellants.

Your argument is like saying: “There’s no way F-15, a plane from the 70s could be faster than F-35, a plane from 2010s”