You could use the picture of the Fakour mounted on the Tomcat and compare length as a ratio to the tomcat’s length between Fakour and Phoenix
Interestingly enough, the diameter of the Improved Orion (M112) is 356mm:
Which matches the diameter of the AD-40A:
So it’s possible that the diameter of the Fakour-90 (without taking that conduit into account) is also 356mm.
@MiG_23M
Iranian SM-1 “clone” (Shalamcheh 2) also uses M112 motor:
And its export version (AD-40) which also uses M112 motor (the picture in the brochure is flipped), also has a diameter of 356mm.
https://mindexcenter.ir/sites/default/files/2022-07/ad40.pdf
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Actually, I think you are right.
AD-40A is also using the same motor (but just the sustainer pour) and warhead.
There is little reason why Fakour-90 would be longer than AD-40A.
Especially for the diameter, we see that the warhead and motor are ubiquitous and even Iranian SM-1 “clone” (Shalamcheh 2 / AD-40) uses it, and yet the diameter is still 356mm (same as AD-40A) and same as Improved Orion.
Fakour-90 wrong dimensions:
https://community.gaijin.net/issues/p/warthunder/i/62aqRWP8TtYp
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It sounds like he calls “Fox 1” when he fires
I had sent an email to NASA and asked whether those figures in the handbook are average thrust or peak thrust.
The question was forwarded to people in the sounding rocket office internally and the answer was: “We have those as average thrust for the motor burn phases.”
Three possibilities:
1- As the English text on the bottom suggests, that’s the AD-40A and it uses or is at least offered with SARH seeker.
2- Some say Fakour 1 is SARH and Fakour 2 is ARH, but it’s not clear if AIM-F90-B is the Fakour 2 or simply a Fakour 1 with a different rocket motor.
3- That could be (an early development) test of firing it in SARH mode. Keep in mind that even AIM-54 can be fired in SARH mode.
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I don’t see the fuel tanks in game, did they add it?
IRIAF never used drop tanks for their F-14s
Inquire how the values were found, is it test data or calculation.
It can be either AP / Poly or ~295 isp but not both.
It does not matter. The fact that the 2023 edition of this handbook has revised the values for thrust and burn time of both phases as compared to the 2005 edition of the same handbook, says a lot. (previous values were incorrect)
We don’t need you to validate NASA’s sounding rocket department’s methods and techniques.
They insist that it’s the average thrust, then it’s the average thrust.
And this is also confirmed by the other study that I’ve mentioned above (section C in the main report) that provides average thrust for a particular test launch.
There is a reason that this rocket motor is so popular for space applications. It’s because of its high specific impulse.
That conclusion is nothing more than an assumption with no foundation.
According to what, real data or calculations? How did they find the number? NASA prides themselves in being able to support their claims most of the time.
Which if this is true, the claimed propellant type is wrong.
All I’m asking you to do is verify what you’re claiming. They responded once already… just ask for clarification but no… you always stop at the most vaguely convenient answer possible because you’re afraid it will harm your narrative.
We have patents from the manufacturer that claim it can reach such high specific impulses even in the lab, let alone in a climb where the continuous decrease in air pressure increases the specific impulse.
And we have detailed information from that particular study in the section C of the main report that confirms that. And shows that even much higher specific impulses than what was specified in the revised 2023 edition of the NASA rockets user handbook or even in the manufacturer’s patent, can be and was achieved, at higher altitudes.
It is actually fine, because it is test data.
These are not explicitly discussing the propellant type used in the I-HAWK, rather the theoretical capability of the type of propellants. These were never realized for stability and longevity reasons as we know.
Such as?
Assuming it is operating constantly at 100% efficiency (never the case IRL) I presume?
Why not email them again and confirm instead of bickering and sticking with the absurdly vague response that agrees with your agenda?
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And?
You claimed that the propellant “type” used in the M112 can’t reach such high specific impulses, and I proved otherwise.
No, in fact they were realized and we know that because NASA’s rockets user handbook as well as that study I provided in the section C of the main report ( https://community.gaijin.net/issues/p/warthunder/i/EaLuXWmNnW4F ) provide us with average thrust and burn time values.
Nice projection.
I’m not the one who’s “bickering and sticking with the absurdly vague response that agrees with my agenda”.
I provided valid sources. (And even went through the trouble to ask NASA’s sounding rockets department to confirm it)
If you “can’t believe them” you should provide evidence to the contrary.
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This is essentially how the missile should look like, if and once they fix all the values (thrust, burn time, propellant mass, dimensions):
No, you didn’t.
Again no, they did not disclose how they came up with the average thrust. You’re assuming things and refuse to inquire further about them because you fear you may be wrong. You won’t even chance devil’s advocate with yourself because you are stuck in a confirmation bias loop.
…
Roflmao
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If you think the NASA sources are correct then it’s completely pointless to keep arguing with him. In fact the best thing would be to just ignore him. I read he stated his "maths"are valid and useful while also stating that atmospheric drag is "near zero"at 20km despite atmospheric reentry drag beginning at around 120 km. I would warrant he’s never even done a line integral once in his life.
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uhhhh i think the drag at ~mach 25 is gonna have a much bigger effect than at ~mach 4-5
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