So do other sources, with the known peak thrust we can easily ascertain the low and extrapolate the rest through the curve.
You can still derive the thrust especially from the REXUS-2 launch seeing as it only has one motor and we know the conditions, burn time, etc.
I never assumed the drag was zero at high speed, I said the drag at such altitudes is already near zero or a non-factor which is absolutely true. The vehicle even without the nosecone and when deploying payload was losing almost no excess energy beyond the pull of gravity.
I am using two different launches, not just one. Quit dismissing additional datapoints because they don’t suit your agenda.
It’s not telling at all, you’re just putting on rose colored glasses and fingers in your ears all the while singing the “lalala” tune.
@Gunjob You wanna weigh in on this? Why not use all available data in favor of the less detailed NASA claims?
Also are you able to forward the following information to the devs? @sudo__su 's 2023 NASA source states “approximately” 20k pounds thrust for 6 seconds and then “approximately” 4k pounds thrust until burnout. This is hardly useful data.
However, if we are to use those as peak thrust datapoints and reference the acceleration curve of the motor based on the following two sources we get a far more complete picture;
The NASA handbook, unlike the study you are quoting, is not providing the data from the last / single launch that they have done.
It’s giving values that are typical and representative of this type of motor.
So this whole “Atmospheric conditions were different in 2005” argument is irrelevant.
Unlike that European study, we are not talking about a single launch with NASA handbook.
Then please calculate the drag for us.
What data?
Rexus 2 vehicle uses a single motor, where the M112 motor is burning from the ground at low altitude, and it only provides you with a single number: peak acceleration at 20g.
ROTEX-T study is using two rocket motors in tandem where the M112 part is burning as second stage at high speed and high altitude.
The aerodynamic / body design of the vehicles are completely different.
Not to mention that the REXUS-2 specifies the propellant mass at 290kg instead of 294kg potentially indicating a less-than-nominal/typical propellant mass.
That means one of your two NASA sources are not credible, the other one also likely incorrectly calculated the performance. The burn time should be considered based on the acceleration curve showing how long the motor is accelerating the missile. From there the thrust needs to be extrapolated based on the given thrust values.
The thrust values given by NASA are between 19,000 and 20,000 pounds for boost and 3,000 - 4,000 pounds for the sustainer.
It’s just that the 2023 revision of that same handbook has revised the values, as compared to the 2005 revision of that same handbook.
Obviously when someone revises something in later revisions, it means that they consider the previous values to be incorrect and the current/new values to be more accurate.
It’s two iterations of the same source, two sources. They changed something. Neither claims match any other available sources and we now know that one of the two iterations was erroneous. Who is to say the newest iteration is any more accurate when they did not cite the source of the information?
That leaves us with the sounding rocket sources I gathered which have shown us the correct burn time and thrust curve.
Your newer “more accurate” data is where NASA says “approximately…” and lists a generic singular thrust value and burn time that doesn’t match any other data?
Exactly my point.
So this is the most accurate up to date information.
And NASA is not using just single launch with some potentially underfilled motor with less than nominal/typical propellant like those European studies …
Doesn’t need to.
No, we know that the first revision was erroneous. That’s why they revised it in the 2023 edition.
I think it’s pretty safe to assume that the source is NASA.
I’ve head they are in the business of launching rockets into space :)
Yeah, single launches that either only provide peak acceleration with a less than nominal/typical propellant mass motor and the other is using it as second stage booster on a vehicle with unknown drag coefficient (And it’s also still a single launch … So just one sample).
I think we can interpret the “approximately” as “average” thrust here.
In fact, If you look at various motors listed in NASA rockets user handbook, you will see that they never mention “peak thrust” for any of them.
They either specify “average thrust” or “approximate thrust”, but never “peak thrust”.
Why?
Because peak thrust is not very useful. What is useful to the audience of the NASA rockets user handbook (engineers and scientists looking to choose a motor) is average thrust. Peak thrust is useless to them …
So essentially, “approximate thrust” means “approximate average thrust”.
It is the least detailed of the credible sources. It provides ambiguous thrust and burn time data. The other source provides accurate burn time and acceleration for rocket of given type. The data provided is sufficient to put the performance in-game within ballpark reason.
It does.
The first iteration shows that they are using a simulation to calculate the thrust whereas the other UK source shows the actual thrust curve and burn time. Using the fact that it peaks at 20G acceleration for one of the missiles during the M112 burn phase as a datapoint we can show that the peak thrust is less than 20,000 pounds with absolute certainty and that the average MUST be less regardless of what realistic % you think the drag has on the overall net force equation.
Peak acceleration of a missile with known mass and with altitude over time as well as a peak apoapsis. This gives us necessary information to ascertain the correct performance.
Acceleration over time shows the correct burn time and thrust curve which proves the 2023 Iteration of the NASA source is not accurate or in-line with actual test data. The previous iteration was also using a flawed sim… no source for the data on the new one.
Then you’re wrong. Simple as.
If they’re approximating an average boost thrust of 19,000 pounds over 6 seconds and 3,000 pounds over 19 seconds you’re talking about 171,000 lb-s total impulse. This is a massive increase over the Phoenix and it doesn’t make any sense.
This is a 71% increase over the Phoenix with only a 42% increase in propellant mass. The math doesn’t math. The propellants have equal efficiency in real life of around ~260s impulse. It literally cannot be the average thrust and I’m using the lower values.
Using the 20,000 lbf and 4,000 lbf thrust values skews it even more. You don’t double the performance of the motor with equal impulse with only a 42% increase in overall propellant mass.
The AIM-54A in-game for reference on the left. The in-game Fakour-90 as of right now on the far right.
The AIM-F90-1 is my test model using the calculated average thrust not accounting for drag. This thrust values are likely too low.
The AIM-F90-2 is assuming 19,000 pounds thrust for the boost stage and 3,000 pounds thrust for the sustainer. As you can see, even if I adjust the propellant mass burned to give the booster a reasonable specific impulse of around 258s the sustainer is absolutely insane and not physically possible with a impulse rating of well over 400. This has not yet been accomplished by any modern propellant.
What is more interesting is that if you use the 19,000 pound average thrust value and the 3,000 pound average thrust value with the 5 second boost, 21 second sustain burn times and adjust the fuel fractions you can actually get a VERY realistic 260s impulse for the booster and ~230s impulse for the sustainer which is what Gaijin already did.
Lacking the actual burn time data and utilizing the 5s boost, 21 seconds sustained as a datapoint they adjusted the thrust values to match expected impulse of the AIM-7F/M propellant type which is produced in the same building as the I-HAWK.
Their thrust values are a little lower and thus the propellant mass differs slightly but with the additional sources they may be able to correct the propellant mass & also the thrust / burn times. Boost phase should last approximately 5-6s with slightly less thrust than it currently has and the sustainer’s average thrust should be considerably less, lasting for about 19-21 seconds.
You are assuming Phoenix has accurate thrust values in the game.
Anyways, instead of this endless argument, if you will, please calculate the peak force/thrust on the vehicle (at peak acceleration) for ROTEX-T and Rexus 2 (separately, for each, as they are different launch vehicles).
Just do a clear snd concise calculation and let’s see what values you will get, for peak force/thrust.
Keep in mind that you should use the mass of each vehicle at the time when it archives its peak acceleration.
No, that’s not what I said or what I did. Stop pretending I said stuff I didn’t to push your side of the argument. This is bad discussion practice.
You’re on the wrong side of the fence and don’t want to admit it so please do this thing you’ve been telling me is useless this whole time?
Using the peak thrust values of 19,000 lb-f and 3,000 lb-f and looking at the % difference between the minimum and peak values based on the acceleration chart we find the sustainer gains 15% acceleration over the duration of it’s progressive burn.
(3,000 * 0.85) + 3000, divide by 2 = 2775 lb-f or 12343.81 newtons. This is roughly how Gaijin determined their 12,400 newton sustainer thrust for the motor already.
Likewise, the booster… less of a difference between minimum full and peak thrust.
19,000 lb-f is 84516.21 newtons. The in-game missile has 83,900 newtons. This is about a percentage difference since the booster is neutral burn and should be a plateau.
Given these fractions and the 5s/21s burn times we find a reasonable output of impulse and overall deltaV which is almost in-line with what is in-game already.
This isn’t perfect obviously and the dev team will ascertain their own values from the available data but telling us to dismiss some of it because it doesn’t suit you just isn’t good reporting practice.
The specific impulse for AIM-54A (In-game values):
430500 / 9.8 / 170.55 = 257.6 s
Specific impulse for M112 (With 2023 NASA handbook thrust values):
854058 / 9.8 / 294.835 = 295.6 s
Didn’t you previous suggest that “considering that M112 is manufactured by the same company and facility and tooling that manufactures AIM-7F they probably have similar specific impulse”?
In the game AIM-7F has an Isp of 328.23 s for the booster stage and an Isp of 301.49 for the sustainer stage.
AIM-7F’s total specific impulse:
(4.5 * 26940 + 11 * 6340) / 9.8 / 61.23 = 318.3 s
So even if we set M112’s thrust values to the values provided by 2023 NASA rocket user handbook (and accept them as average thrust), the specific impulse for M112 will still be worse than AIM-7F’s motor and right in between AIM-7F and AIM-54.
Difference in specific impulse for M112 (2023 NASA handbook values) and AIM-54 (in-game values):
295.6 - 257.6 = 38
Difference in specific impulse for AIM-7F (in-game values) and M112 (2023 NASA handbook values):
318.3 - 295.6 = 22.7
So the 2023 NASA Handbook values actually are quite reasonable and are most definitely “approximate” average thrust values (just as every other thrust value for other motors listed in that document are average thrust values … You won’t find any peak thrust value in that document, because it’s useless to its audience).
I haven’t been following too closely along, but is the in-game Fakour’'s ISP closer to the thrust values of the AIM-7F motor irl or the thrust values to what is in-game? Because we know the AIM-7F has unrealistic thrust values in-game to account for the longer burntime. Did Gaijin do the same with the Fakour?
The AIM-7F thrust in-game is incorrect, propellant weights are swapped incidentally by Gaijin from the standard characteristics page but it has not been fixed. This resulted in an overperformance of the motor with 37,000 lb-s total impulse instead of ~30,000.
No, an ISP of 290 was not possible until VERY recently. This is common knowledge for propellant discussions. The ISP of 260 for the AIM-7F was breakthrough at the time and was shortly followed by a switch to HTPB from CTPB types as HTPB overtook the performance in specific impulse around the mid 80s and early 90s. It was not until 2010+ that 290+ ISP motors were in true mass production at the earliest and we’re discussing a missile that went into production in 1971 - I’ll consider that a notion that it was likely a little bit less than 260s impulse but I gave it the benefit of the doubt anyway when adjusting the data.
That isn’t why it has additional thrust afaik. That was just an assumption from the community but the propellant masses are wrong so the impulse calculation is far off.
Even with a Isp of 240 (worse than AIM-54A) what we have in the game is still underperforming.
Plus, it’s not about whether the Isp values of the AIM-7F and AIM-54 in the game are realistic.
If they had to give these missiles higher thrust due to their long burn time and “because a missile’s drag is reduced when it’s firing and we don’t have that modeled in the game engine” then the same would apply to M112 as well …
170.55 / 294.835 = 0.5784 - 1 = -0.4215
That’s how I came up with the numbers, but yes you are correct. It would be 58% of the mass of the M112’s propellant or be lacking 42% of the mass of the M112.
Interestingly, the switch to HTPB offered better longevity in storage for rocket motors over older CTPB based propellants.
For very long range missiles that loft they don’t need to do this, they spend most of their time at higher altitudes where drag is considerably less of a factor. Gaijin’s drag model at altitude is erroneous either way so they need to adjust down the drag value of the missile. Like the AIM-54A, this will probably cause overperformance at lower altitudes already even without the additional thrust to overcome reduction in drag during motor burn time.
This additional thrust is not modeled for the R-27 series, AIM-54 series, or any other top tier missile and was only an assumption by us that it was done for the AIM-7F.
