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:
1-
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.
3-
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”).
5-
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.
