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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Yeah!
this only goes up to mach 1 where drag is a lot higher than if you were going faster
" The value of the drag-divergence Mach number is typically greater than 0.6; therefore it is a transonic effect. The drag-divergence Mach number is usually close to, and always greater than, the critical Mach number. Generally, the drag coefficient peaks at Mach 1.0 and begins to decrease again after the transition into the supersonic regime above approximately Mach 1.2."
Drag-divergence Mach number - Wikipedia.
yes its Wikipedia but I’m just a dumbass that knows a bunch of random things
Drag is proportional to velocity squared
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You can do the maths yourself. I used the frontal area of an AIM-54 with the drag coefficient of a G7 projectile at 1475 m/s with an air density of .092 kg/m^3 at 18,000 meters. I got 22107 Newtons. If you try the same at 7800 m/s at 120,000 meters I got 4978 Newtons. It’s around 4 times more drag at 15 percent the altitude and 19 percent the speed.
The fact that the guy @sudo_su1 is arguing with honestly said that 44 m/s^2 is “near zero” for his “maths” to be considered correct is hilarious.
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The math is still correct and valid, the resultant number just hasn’t been corrected for drag. You guys saying it’s invalid doesn’t make it invalid - if math is invalid it would be something you could show… mathematically.
Now if we could all hop off the insult train and focus on proving / disproving the motor has 290+ seconds specific impulse we’d be getting somewhere. They’re currently waiting on @sudo_su1 to email NASA back and ask them how they derived the average thrust and what data was used.
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AIM-23C Sedjeel ground separation tests (In the 80s during Iran-Iraq war)
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After a lot of digging, this might be the test firing footage of an actual Fakour missile with boost-sustain motor (rather than the sustainer-only AD-40A export variant/offering, footage of which often circulates as “Fakour test launch” where the pilot uses the “Fox 1” code during missile launch)
This is sustainer-only AD-40A export variant/offering for comparison:
And this is the AIM-54A for comparison:
The perspectives are all a bit different, but I think the difference in acceleration is still obvious.
@MiG_23M I’ve sent a bunch of emails (To both NASA and Aerojet’s parent company L3Harris). I will let you know if they answer.
Safe to say that they launch a bunch of these things:
Also also:
I’m pretty sure he’s taking drag into account though ;)
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There is also a 2015 edition of the “NASA sounding rockets user handbook” and it provides different figures than both 2005 and 2023 editions.
2005 edition:
19000 lb @ 4s + 3000 lb @ 21s
2015 edition:
20000 lb @ 6s + 4000 lb @ 19s
2023 edition:
20000 lb @ 6s + 4000 lb @ 18s
So as you can see, this is a very high quality source that every few years, with each new edition, refines and revises the thrust figures that it provides, based on the expanded data set and improvements in measurement and calculation accuracy (of velocity and altitude and drag etc).
And that’s besides the fact that this is also the ONLY publicly available source that gives thrust figures for this motor. (Other than the test launch study that I linked in the section C of the main post in the bug report that gives average thrust for the whole burn time)
(Regarding this report: https://community.gaijin.net/issues/p/warthunder/i/EaLuXWmNnW4F )