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?
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.
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
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.
uhhhh i think the drag at ~mach 25 is gonna have a much bigger effect than at ~mach 4-5
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
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.
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.
AIM-23C Sedjeel ground separation tests (In the 80s during Iran-Iraq war)
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 ;)
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 )
This just looks like the latest launch data was used and the performance of the motor varies quite a bit (as expected from that older propellant).
As you know, the other sources I initially shared with acceleration over time shows the thrust indirectly. Gaijin has been known to use this as additional source material for confirming other datapoints in the past and they should definitely look into it here as well.
As I just showed here, they launch many of these each year, so they have a very large sample size. Way larger than that one singular European launch that you are trying to use to guesstimate thrust from, without even taking drag into account or knowing the vehicle’s mass at each point in time.
And no, old propellant doesn’t cause a 38% increase in total impulse. That’s just not how rockets work.
Less accurate measurements or underestimation of drag does …
As I and others have pointed out already, that source is virtually useless for calculating thrust.
Plus, one singular launch is essentially noise compared to NASA’s data which is based on many many launches.
Not only is expired propellant a possible reason - it is the most likely reason the total impulse changed so drastically.
There are few other reasons explaining the discrepancy with such high likeliness. I already shared a document showing that certain batches of the motor had unbonding issues.
You and others can remain incorrect
That is your assumption, that they base the date on many launches. I would think that with such a large datapool as early as 2006 that the values would not change so drastically year to year unless the tolerances and performance of the motor varies so wildly.
If the values from the 2005 edition of the handbook (which result in a lower total impulse) were due to poorly aged motor it’s all the more reason to use the newer, revised and more refined values from the 2023 edition of the handbook that are based on the expanded data set / larger sample size.
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.
