tavarish palkovnik

Sorry @GGTharos if I bother too much but you didn’t fully convince me just with this statement. I would need more Simply I don’t see this motor as single thrusted, even this only one available crosscut photograph, although blurred, doesn’t implicate single thrust…as per my opinion Momentarily 80% dual thrust and 20% on some ultimate single thrust

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You didn't help much but thanks... what is IRL F-14 weapons manual? By the way, I'm also very interested in history as well and would like to have some answers how ones took from others and opposite. This motor is just one beside others where I can see so much similarity

Until I see geometry of grain which will allow continuous thrust of 15000+ N in time of 20+ seconds, for me all options are on table Which source is the most reliable that firmly state singularity of thrust?

After some time to write again something about AIM-54 This is only available crosscut view of motor, AIM-54A or AIM-54C I don't know, but actually and later about it, based on my opinion it is not so important. Although picture is very blurred it seems like fuel block (fuel grain) is with cylindrical inner hole and outer backside surface in 1/3 of length is exposed while outer front side is inhibited. This is how it could be where with red lines is marked initial burning surface Some years ago, I was reading some old soviet document (which unfortunately I can't find anymore) with description of NATO weapons systems, and for AIM-54 it was stated that motor is dual thrust ! This configuration if it is like that, indeed gives two thrust sequences, simply burning surfaces are such Here I made some calculation based on this geometry and with estimated characteristics for let's say Mk47 pk AIM-54.pdf First half slightly progressive buster stage, second half slightly degressive sustain stage with total working time of 24 seconds. By the way, volume and geometry allow some 170 kg of fuel what is in line with some documents. And for sure, either Mk47 or Mk60, I'm giving to both nearly same weight of fuel. When about those two, with keeping same geometry and weight and just with changing fuel, keeping same energetic (chemical) properties but just with different burning law, more or less same function form will be in both cases, with slow burning fuel and with a bit faster burning fuel. Total impulse is slightly greater with faster burning fuel, 5% difference I am not claiming anything, just giving my opinion, for me and I believe that, motors of AIM-54 are with dual thrust concept

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It could be 47800 kgs at sea level what is significantly different to what you are giving for R-33. Although these two motors are from two different design bureaus, older one is quite enough known, in geometrical form of fuel block and burning time, there shouldn’t be some drastic difference between those two. Last summer one patent is published, and as one of authors signed is personally V.A.Sorokin, general director of MKB Iskra, and Iskra designed and manufacture this motor. Description and model sketch for me is nothing but this motor. Eventually it could be some redesign of motor of H-58 (originally also their) but even if it is this, still to me all these R-33, R-37 and H-58 motors (380mm) are more or less very familiar. Familiar in geometry and fuel mass.

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On few available graphs where relation between thrust forces at SL and altitude is given, it can be seen that differential is not constant. Of course I understand this 7% is unification but just to be mentioned. Increase is more rapid between 0 and 10km because function p f(H) is such as well, on higher altitudes atmosphere pressure values are more equalized. And this 7% is fine but it is important to place each individual motor on it's place in this graph. As we can see, there are motors with expansion under 1 bar and also over one 1 bar This is one more nice motor in 380mm, well known Kh-29 Just from curiosity, what could be pressure at nozzle exit on this motor. This equatation is very useful, with chamber pressure and ratio of areas (Exit vs throat) ''pe'' will show it's value. To get ''pk'' and ''A'' there should be some work but without work nothing can be done...(F tot 23300 kg, t=4,5s, mg=110kg, probably VIK-2, Isp 225s (40:1), Ief=207 T=2820K, k=1,2, dk=80mm, de=210mm, pk=6,5MPa, A=7 ... etc ... etc ... -> pe=0,13 Mpa So it can be said, this motor, this rocket, is with thrust of 50800 N at SL and at 5km it could be 53430 N or +5%, at 10km 54380 N or +7% etc

Don't mind to discuss it at all. Drag force ''Fd'' = Drag coefficient ''Cx'' * dynamic pressure ''q'' * Referent area ''Sref'' As you can see, I always try to specify together with Cx what is referent area for calculated drag coefficients, and mostly it is body cross section area instead of fixed values like square meter or something else. 0,8 for Sref 0,05307m2 is roughly 0,0424 for Sref 1m2. Roughly because it doesn't go that way exactly

Some mathematics about this... So based on this, I'm getting that this motor is optimized for altitude of 10km with pressure at nozzle exit of 26 kPa. With same principles motor of Neva 5V27 is with pe=80 kPa or optimized at altitude of 2km

More or less but mostly everything can be calculated with just simple involving elementary study of rocket motors. Luckily there are plenty of motors sharing same diameter 380mm or 15’’ and having similar composite fuels so even visible differences forcing directions. All these three are in 380mm 5V27 of S-125 Neva (SA-3), R-33 and Kh-58…all in same diameter as AIM-54 5V27 is primarily for low and near to intermediate altitudes and nozzle is sized respectively Kh-58 is more for low intermediate and near to low altitudes This motor’s cross section is my creation but I stand for it very firmly And this is Mk47 Mod 0 with nozzle bell sized, obviously not optimized for sea level or intermediate altitudes but for up there altitudes just as R-33

For start one fun fact…commercial jets which usually (from clear technical reasons) fly at 11km at 0,8M where temperature is roughly -60degC don’t have artificial heating of wings where fuel tanks are located because air friction energy do that job instead. On burning rate of solid rocket fuels, primarily pressure in rocket motor’s chamber influence. Temperature of fuel grain also. But pressure primarily. Different temperatures and accordingly variable pressure will make one differential, different pressures and accordingly variable temperature other differential. This other is significantly bigger. Examples…this is 9M330 motor of TOR, but from time when 9M330 of that time was in form as it was. Later 9M330 and it’s characteristics is changed but what I want to show is not changed. As you can see, total impulse (суммарный импульс as written, for those who struggle with Russian language) is quite close at +50 ; +15 and -50degC Another one, 5V55 of S-300 тяга двигателя -> motor thrust and this is in k(1000) of kilos. Same story. And third sample is archaic R-3… All right, temperature made influence. Total impulses should be roughly in line but altitudes influence also. Not on burning rate actually but like altitude does, with pressure. And by the way, this motor, this nozzle has very poor expansion compared to some other “church bells”

Base drag is something else and actually not related to this topic which is more about internal ballistic. Here I made some simple text about base drag or better to say, about difference of drag coefficients in active and passive stages if someone finds interesting -> But to return on equation, it is empirical equation in form Ical=Th*Inom+190,3+76*pk-3,058*pk^2-7000*pa+25484*pa^2 pk- chamber pressure (MPa) pa- ambient pressure (MPa) Inom- nominal impulse, theoretical impulse, maximal impulse. Fuels of USA origin are usually given in ratio of 70:1 while of Russian origin in ratio 40:1 (40 bars in chamber vs 1 bar ambient) And Th are coefficients of losses which are always present. I took some maximal values, losses of unburned fuel, losses of friction in nozzle and losses due to effect of condensation phases in exhaust

I will later today...in the meantime just to make one correction, it was with typing mistake -> I=245*9,81*0,96*0,97*0,98+190,3+76*5-3,058*5^2-7000*0,1+25484*0,1^2=2242

This is how it looks in my ''shots'', AIM-54C ballistic flight, 45000ft, 1,2M, 0;30;45 deg, constant thrust of 13595N in 27 seconds But, thrust of 13595 N is something about what I would like to talk. 13595*27/163=2252 kg*s/kg ... it means impuls is 2252 or 229,5 s what seams is quite less then regulary used 250 or 245 or 260 etc One example for start, rocket motor of Neva 5V27 rocket Fuel in this motor is known, composite with maximal theoretical impulse of 245s at pressure ratio pk:pa=40:1 Diagram says 345300/151,4=2280 kg*s/kg or 232 s what is 5% less then theoretical maximum at 40:1. It can be seen that average pressure in chamber is 50 bar and test is performed at pa=1 bar. So mathematic will say for this -> I=245*9,81*0,96*0,97*0,98+190,3+76*5-3,058*5^2-7000*0,1+25484*0,1^2=2242 or 228,5 s In practise, motor with fuel with maximal impulse 245s at 40:1 will give real 228,5 s at 50:1 with combined losses in burning process. What about AIM-54C and these ''shots'' with 13595N of thrust. I have one fuel composition which I believe is or it is close to what is inside. It is fuel with maximal theoretical impulse of 247 s at ratio pk:pa=70:1, density is 1,62 kg/dm^3 and burning rate at 70 bar is 5,3 mm/s. Let's say nozzle throat is 55mm, chamber pressure should be around 4 bar -> 13595/1,4*4/3,14/55^2 I=247*9,81*0,96*0,97*0,98+190,3+76*4-3,058*4^2-7000*0,1+25484*0,1^2=2211 or 225 s So, this motor in practise at sea level will have thrust as result of impulse of cca 225-230s. But these shots are up there at high. Configuration of nozzle is such to allow expansion of gases to the pressure quite less then 1 bar, based on some calculations I'm getting 25000 Pa what is in line of ambient pressure up there. So what exactly should be impulse of motor burning there -> I=247*9,81*0,96*0,97*0,98+190,3+76*4-3,058*4^2-7000*0,025*25484*0,025^2=2498 or 255 s 2498*163/27=15080 N or 10% more then at sea level Higher altitude -> higher impuls Higher altitude -> higher Cx In some cases what impuls gives drag takes in roughly same percentage but here taking in consideration nozzle exit and expansion, I belive impuls will give a little more then what drag will take and by that, mach numbers should be more then as presented at the beginning, by my opinion...unless flight is not ballistic but with kinematic overload

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I have something else in addition to ask. This could be more tougher. This is the only one cutaway model that can be found on the net. I’m the most interested in motors and it would be fantastic to be able to see it in real in close look…but…not possible So what I wanted to ask, did anyone perhaps visited this museum and have pictures not blurred like this. This motor could be MK60, seems like cylindrical grain inhibited on outside in 2/3 length, burning from inside, from up and back heads and from outside on uninhibited surface. From geometry it should be roughly 200kg of fuel and with some very regular burning rate of 7mm/s that gives 20 seconds of active

Thanks The_Tau, it has sense and it is more correct than many fairytales that can be found on Internet

Can anyone help with translating this?

Today for fun I modeled this flight, for fun and relaxation. Are we close with other variables of this flight? Just from curiosity

Thank you Маэстро, understood and accepted. Although for physics and mathematics all times are good, I (we) will leave it till some better time comes