What Are Our Motors "Missing"?

[Adrian A]

very cool stuff, and thread. i wish there were more topics on tailoring rocket design to achieve utmost efficiency. i imagine even basic commercial F and G motors could be attaining much more altitude with some tweaking in the airframe design.

Here are a few examples. In case you need something to help you get to sleep:

https://www.rocketryforum.com/showthread.php?t=26383
https://www.rocketryplanet.com/forums/showthread.php?p=91085
https://www.rocketryplanet.com/forums/showthread.php?t=1999
https://www.rocketryplanet.com/forums/showthread.php?t=2808
https://www.rocketryplanet.com/forums/showthread.php?t=4081
https://www.rocketryplanet.com/forums/showthread.php?p=134610
https://www.rocketryplanet.com/forums/showthread.php?t=5667

 

[mikec]

it does beg the question, especially of EXers, why has nobody really done a dart attempt at 200k+ AGL.

RRS flew a boosted dart to about 50 miles in 1996:

https://www.rasaero.com/dloads/RRS 50 Statute Mile Boosted Dart.pdf

 

[deandome]

Daveyfire, et al...thanks for your patience!

I do much better with simple explanations like "it's the insulation" and "...cat hair" & tend to gloss over the equations and engineerese.:blush:

And I'm not a troll...I just haven't been active in the hobby the past year or so, and I usually read @ Rocketry Planet vs. perusing threads.

Over the winter, I'm going to get all my older stuff in flying shape (some damage to my Thor had me unable to fly my larger motors) & build something new from the parts I've been sitting on for like 3 years now.

Like I said initially, in looking at missiles for ideas, I got really curious as to how that small/light Arcas could hit 32 miles when guys with access to much newer/better materials & tech are spending $10K or so launch much bigger/heavier rockets that hit 1/2 that altitude (Gene Nowakcyk). You might say people answered it earlier in the thread, but they really didn't explain why it has not been done (or even come close, for something that size) the hobby level until the last few posts.

Still seems like it was a good question for this board, at least to me. As to a solution, for a single-stage 64 lb rocket, I'd think you'd have to build the entire rocket as "single use" (still with recovery though...and it might be reusable)...lining some pipe with a thin layer of high-tech insulating goop & pouring propellant into it.

I'm sure these guys have some scrap 4" pipe lying around that they'd give to a hobbyist for a song :wink: https://www.ticotitanium.com/products/titanium-pipe-2/titanium-seamless-pipe/

Whaddya think a 6' segment of the seamless lightweight pipe would be? :y: Maybe the welded pipe would be cheaper, and/or it could have thinner walls, if that could survive the long burn.

 

[Reinhard]

An M350 would be cool, but not many people could either provide the light weight and tower necessary to get it going fast enough at liftoff, or better yet, the kick-butt booster you'd need to get it going really straight.

AT is already halfway there, with the L339. I guess even this one is more of an halo product than a great seller.

As I think about the ARCAS, the breach loading launcher is probably the biggest show-stopper about getting similar performance from a single-stage hobby motor. But put a small Vmax or Warp-9 booster stage under it to get it some speed out of the tower, and now you're talking.

Launching a rocket through a piece of Sonotube doesn't sound that hard to me. Of course, some attention has to be paid to details like the volumes/pressures and the attachment/detachment of the sabot to/from the rocket. The result may not be as impressive as the ARCAS, but maybe worth the effort. Transporting the launcher seems like a pain to me. My biggest concern is of a legal nature. I assume this type of launcher comes close enough to the legal definition of a cannon to cause some trouble or at least unwanted attention.

Reinhard

 

[JPVegh]

At the risk of feeding the troll I have to say that this was one of the most interesting threads I've read here in a while. Thanks Deandome for asking a simple question that spurred the conversation.

 

[Adrian A]

I don't ever plan to mix my own propellant, but this has me curious about a couple of things, because I like a lot of the advantages of end-burners.

Are there pourable variations of propellants that would be worthwhile, performance-wise, in the interest of more easily embedding wires?

When we won the BATFE lawsuit, it was because we could prove that APCP doesn't burn fast enough to qualify as an explosive material. How close is Warp-9 or Vmax propellant to that threshold? If I had an endburner that burned at, say, 4 inches per second to give me a 15-second burn with a 60-inch long motor, is that over the line? If wires are embedded to increase the burn rate, would that count against us for BATFE burn rate purposes?

Are there reasonable ways to do with the liner erosion issue? Could you bond in another liner on the inside for the bottom half of the motor?

 

[ClayD]

Adrian... Making propellant is fun! (you and I both have reasons for extreame cautions if you ever decide to dabble!) But, I do have a scrap burning buddy...

There are some people that use double liners for long burning motors. Namely the offset core, has a high mass flow down the side of the case later into the burn. (or the point the propellant on the thin side of the core breaches to the liner) I have never seen anyone only use the double liner on the bottom half.. (But i really like that idea...)


Your BATFE comments are a fairly limited scope of what occured for burn rates. It also was not why we won. "a capricious and unlawful abuse of power" I find to be the main reasoning in Judge Waltons letter. Burn rates were symantics. In making an amature propellant, the BATFE is not going to be tooo considered of for burn rates. The reasoning here, is something not for a forum.. The FAA only cares about impulse and time... I thought the burn was limited to 55 seconds? and thats for trajectory reasons.

Since your not buying or selling propellant, theres no reason to discuss the batfe "want" to require an LEUP for motors in commerce as relates to burn rates.

If I had an endburner that burned at, say, 4 inches per second to give me a 15-second burn with a 60-inch long motor, is that over the line?

(How long is an N1000 burn)... so to answer your question... the burn rate is not in motor configuration! (i know you knew that, your paragraph confused me)
that goes to say, and end burner that operates at fairly low pressure, will have a faster sea level pressure burn rate, than does a moon burning propellant that burns in the same time frame. (Pressure exponent aside)
(again, the burn area is not linear in a moon burner, but is on an end burner - so the end burner would have to be several times faster)
I may very well be wrong here, but warp 9, hapens to be a fast propellant, but not much faster than other "fast" propellants at ambient pressures... It has a very high burn rate exponent, which means if presure goes up by 1 , burn rates go up by 5 or 6(*heressey-not real numbers..*)
Thats why you see a J1299, and H999 in .5 second burns, and and end burner of 8 seconds. LIke i said i could be wrong, but thats what i *think*.

you guys have a blessed day!

 

[jsdemar]

Are there pourable variations of propellants that would be worthwhile, performance-wise, in the interest of more easily embedding wires?

Pourable doesn't necessarily make embedding wires easier. Standard processing steps for high-performance propellants are not compatible with having strands of wire in the mix. There are other ways to increase burn rate.

When we won the BATFE lawsuit, it was because we could prove that APCP doesn't burn fast enough to qualify as an explosive material. How close is Warp-9 or Vmax propellant to that threshold? If I had an endburner that burned at, say, 4 inches per second to give me a 15-second burn with a 60-inch long motor, is that over the line? If wires are embedded to increase the burn rate, would that count against us for BATFE burn rate purposes?

If the material allows a transition from flagration to detonation, it's explosive. Burn rate is one factor. Properties of the ingredients is another. And the depth or thickness of the material is also important. Propellants like Warp9 are no longer used by the military because the metalocenes will migrate over time and produce a sensitized material. In larger motors, they fail the detonation tests.

A cursory search of the professional literature on the subject will show the trend for current energetic propellant formulations. It doesn't include embedding wires.

Are there reasonable ways to do with the liner erosion issue? Could you bond in another liner on the inside for the bottom half of the motor?

Yes, or taper the thickness of the liner continuously. If you want the case to be reusable *and* pass the NFPA1125 case temperature requirement, it will not have a high volume loading due to the extra liner thickness. One could make an EX motor where the case temperature is just about at the melting point of aluminum when the motor burns out. But, don't expect your high-temp epoxy airframe to survive the heat soaking during the rest of the flight. You would have to fly the motor with a metal fin can. Now we've jumped the fine line from HPR to amateur rocketry.

 

[Reinhard]

When we won the BATFE lawsuit, it was because we could prove that APCP doesn't burn fast enough to qualify as an explosive material. How close is Warp-9 or Vmax propellant to that threshold? If I had an endburner that burned at, say, 4 inches per second to give me a 15-second burn with a 60-inch long motor, is that over the line? If wires are embedded to increase the burn rate, would that count against us for BATFE burn rate purposes?

As far as I know, the BATFE lawsuit did not provide a threshold or any insight where this threshold might be. Basically APCP was declared an explosive "because we say so". After this was challenged BATFE produced data, but this was scientifically not sound and easily refuted. So basically it was determined that BATFE failed to show that APCP is an explosive, but the question what an explosive is wasn't further explored.
Thats somewhat unsatisfying from a scientific point of view. Someone could contact the BATFE and "ask for clarification".
:duck:


This discussion inspired me to two ideas. They probably have some unknown drawbacks, because I imagine there were people who had them before me.

1) An endburning coaxial motor design with two different propellants. The outer layer consists of an propellant which is optimized for Isp. The core, which can be quite small in diameter, consists of a fast propellant. This one acts as a "pacemaker" for the outer propellant. The burning surface would look similar like in the designs that contain one center wire, a cone pointed to the front of the motor. Compared to the designs using a wire, the advantage would be that the steady burn rate only depends on the speed of the center propellant. The center propellant wouldn't even need to be APCP, as long it is compatible to it and has the desired burn rate. If it were much faster then the outer propellant it would result in a motor that resembles a core burner. This could be fine tuned to maximize volumetric loading without getting erosive with the drawback that the motor takes a little time to "ramp up".

2) Carbon phenolic composites. Theoretically this could be a strong and very heat resistant material for motor casings, even if it is not further pyrolysed to carbon-carbon. Its probably brittle and has manufacturing issues. Phenolic resins don't sound very hobby friendly to me either.

Does anybody know if something like this has been done or, more likely, why it is not feasible?

Reinhard

 

[deandome]

What about plain ceramic liners.? I wouldn't think you'd need anything too exotic. Or high-temp, Pyrex-y glass?

And if they used asbestos/other fibrous stuff as an insulator on the Arcas (see how I bring it back 'round?), what about a thin sleeve/layer of Kevlar cloth?

I'd think if you're willing to sacrifice a bit of power for a given motor diam/length combo in order to have a hot end-burner, you could do some kind of a double-liner (like someone mentioned earlier).

BTW...what's this about a NFPA case temperature limit? Is it based on the melting temp of aluminum (burn thru), the ignition temp of cardboard (motor tube), or something else? Cuz you could make the casing or the tube out of higher-temp stuff if need be.

 

[Adrian A]

This discussion inspired me to two ideas. They probably have some unknown drawbacks, because I imagine there were people who had them before me.

1) An endburning coaxial motor design with two different propellants. The outer layer consists of an propellant which is optimized for Isp. The core, which can be quite small in diameter, consists of a fast propellant. This one acts as a "pacemaker" for the outer propellant. The burning surface would look similar like in the designs that contain one center wire, a cone pointed to the front of the motor. Compared to the designs using a wire, the advantage would be that the steady burn rate only depends on the speed of the center propellant. The center propellant wouldn't even need to be APCP, as long it is compatible to it and has the desired burn rate. If it were much faster then the outer propellant it would result in a motor that resembles a core burner. This could be fine tuned to maximize volumetric loading without getting erosive with the drawback that the motor takes a little time to "ramp up".

I like the idea of the coaxial motor. It seems like the burn would be paced by the faster-burning core, since if the outer ring were slower, then the difference would just expose more area for it to help it along. So that's cool. But we'd still need a core that would burn faster than Warp-9 propellant to get the 4" per second or so that would allow a full 4" O to end-burn over 15-20 seconds.

I like the idea of an end-burner's fairly constant thrust profile. That seems like it would provide the advantage of allowing you to perform the whole burn at near the max design pressure of the case. With a constant thrust profile and a high propellant loading, that would provide increasing acceleration throughout the burn. If that's coupled with a burn that's long enough to get above most of the atmosphere, then you get another benefit of having most of the delta-V happen at a lower density.

But when I look at Cesaroni 4" motors, I was surprised to find that with 6 grains of classic propellant, the moonburner (N1100) was actually a little more efficient, in net Isp, than the same propellant in a bates grain (N2500). I wonder if the moonburner has a smaller nozzle and higher pressure at the start of the burn, compared to the version with the bates grain that has a relatively constant thrust profile.

 

[stickershock23]

Wow this got heated quick.. ( about the end of page one start of page 2)

Look I used to work for someone building HOBBY motors..
they also did other stuff.

We built 7' tall rockets (boosted darts... I will say NO MORE) that were going to the altitudes you say we can't do.

They were using almost EXACTLY the same propellant we can buy for the hobby.

the only difference. How the "grain" was cast, and how it was Checked after it was cast (x-ray etc)

So if (like i said in my first post in this thread) you have enough money, yes you could build a hobby rocket with the same (or at least really similar) propellants we use. PERIOD.

Sorry I will NOT go into what they were, what they were used for etc etc etc. But it is possible

Could I build one myself... If I had the money and felt like wasting it on something that crazy.. yeah I could build one, but I have 0 interest in doing it.

Daveyfire.. I will talk to YOU more about it at MWP if you want..

why won't I talk about it.. because no matter what I say someone will say BS or they will say I am talking bad about someone..

 

[deandome]

Not to heat it up (too much) again, but I wasn't arguing about just hitting those altitudes, I was talking about hitting those altitudes with a 64 lb, 7' tall x 4" diam, single stage super-basic design rocket using about 30-40 lbs of propellant

I'm not arguing any more, but I think you're boosted darts must have been a whole lot bigger/heavier/more complex than that.

Wow this got heated quick.. ( about the end of page one start of page 2)

Look I used to work for someone building HOBBY motors..
they also did other stuff.

We built 7' tall rockets (boosted darts... I will say NO MORE) that were going to the altitudes you say we can't do.

They were using almost EXACTLY the same propellant we can buy for the hobby.

 

[stickershock23]

Not to heat it up (too much) again, but I wasn't arguing about just hitting those altitudes, I was talking about hitting those altitudes with a 64 lb, 7' tall x 4" diam, single stage super-basic design rocket using about 30-40 lbs of propellant

I'm not arguing any more, but I think you're boosted darts must have been a whole lot bigger/heavier/more complex than that.

Dean..

No problem. I did not say anyone was arguing.. I just said it was getting a little heated... and interesting.

Listen to Davefire.. his left pinky is smarter than me. and I consider myself pretty brainy..

I will say one thing about my post.. the rockets i am referring to are NO WHERE NEAR a hobby rocket. they were 100% single use. so you are right.. if you follow hobby rocket rules and regs.. you probably cannot do that kind of altitude... but our propellant is capable of it..

 

[ClayD]

But when I look at Cesaroni 4" motors, I was surprised to find that with 6 grains of classic propellant, the moonburner (N1100) was actually a little more efficient, in net Isp, than the same propellant in a bates grain (N2500). I wonder if the moonburner has a smaller nozzle and higher pressure at the start of the burn, compared to the version with the bates grain that has a relatively constant thrust profile.

Bates style grains typicaly have larger nozzles, since towards the end of the burn your KN ratio can get HIGH.

Moon burners have a realatively constant burn area, so th KN range can be narrower alowing you to narrow in on a nozzle diamter and pressure parameter.

Also if you look at the "startup" burn surface area, a bates has a much higher surface area than does a moon burner. So you KN is a lot lower, allowing you to have a higher pressure.

if your KN is high your pressure must be low, if you kn is low your pressure can be higher.

I am still elementry so theres probably a better way and possibly more accurate way of stating that.

BTW...what's this about a NFPA case temperature limit? Is it based on the melting temp of aluminum (burn thru), the ignition temp of cardboard (motor tube), or something else? Cuz you could make the casing or the tube out of higher-temp stuff if need be.

Dean, we use a particular alloy for our cases, we do not use *steel* aerospace or drill pipe...
NFPA is a chosen and adopted code for TRA research, and TRA commercial codes. Case temperature is based on 6061 aluminum, and at that temperature the strength turns to crap. (its all about maintaining strength)

Composite wound cases are the only other accepted case material i have seen used. I personaly would like to see some research on thier failure modes, to discern if they are more dangerous than is a 6061 case. So if you want better than 6061, you can make a composite case. *you could probably even use thinwall 6061 and composite over it....*

And if they used asbestos/other fibrous stuff as an insulator on the Arcas (see how I bring it back 'round?), what about a thin sleeve/layer of Kevlar cloth?

Google 3M nextel space cloth... they even have a milled fiber than can be used as ablative strands.

 

[Adrian A]

Bates style grains typicaly have larger nozzles, since towards the end of the burn your KN ratio can get HIGH.

Moon burners have a realatively constant burn area, so th KN range can be narrower alowing you to narrow in on a nozzle diamter and pressure parameter.

Also if you look at the "startup" burn surface area, a bates has a much higher surface area than does a moon burner. So you KN is a lot lower, allowing you to have a higher pressure.

if your KN is high your pressure must be low, if you kn is low your pressure can be higher.

Maybe this is the blind leading the blind here, but I thought it worked a little differently. The thrust is dependent on the pressure and the throat area, and I think it's mostly independent of the propellant type. Higher pressure is always more efficient, but it's limited by the case design. You can increase the pressure by increasing the burn rate, increasing the propellant burn area, or decreasing the throat area. The latter two are the Kn. So if the burn area is constant during the burn (as in an end burner) you should have nearly constant thrust and pressure. Bates grains produce nearly constant burn area also, since the cylinders get shorter as the core area increases. Bates grains have a much higher burn area than an end burner, so the throat area has to be larger to prevent the pressure from being too high for the case. Offset-core moonburners have a variable burn area that starts out high and goes down as the burn progresses, which you can see in the characteristic triangular thrust curve. They also have a non-centered mass that can affect the rocket dynamics. Did I get that right?

 

[ClayD]

Maybe this is the blind leading the blind here, but I thought it worked a little differently. The thrust is dependent on the pressure and the throat area, and I think it's mostly independent of the propellant type.

Yes, and no. THeres some very complicated dynamics at work with the combustion and pressure generated.

The heat, density, and other things are all parts of the thrust equation. Pressure and throat area are only part of it.

So the propellant type, can degass to a less dense flow or a more dense flow, changing your thrust resulted.

you see why lox/methane use a lot more stuff to get the same ISP as solid APCP

Slowing the core flow down to cause further degasing is used to create a more ideal flow in some cases.

The pressure in the combustion chamber at any point in time is directly related to the Kn. As the Kn increases, the chamber pressure increases. This relationship is not linear – doubling the Kn does not double the pressure. The burn rate exponent is the fundamental reason for this nonlinearity, typically producing three times the chamber pressure when the Kn is doubled.

I used the term "BATES" loosely i may add. Bates can have ratio 2:1 L/D i think is close to neutral depending on the core size. But a logner grain 4:1 l/d gives you progressive burn KN/Pressure. Can have a wide range of KN.
So you can adjust your core size and your L/D ratio to talor a burn profile.

So if you have 3 times burn rate exponent on the pressure, a narrow KN and a wide KN have really different thrust profiles.

An end burner, will have a fairly Constant KN since the burn area and non-eroding nozzle throat dont change.. But if the exit temperature in a long end burner decreases 5% through the burn you will see a decrease in thrust.
or, if you have more complete degassing as you progress through the burn, increasing your temperatures, lowering the density of the flow .. you will have a change in thrust as well...

This means you could actually see your pressure rise and drop inside the motor not based on the KN, but ballistics of the burn inside the motor.


Talk about blind, I probably murderd that...

*what i was trying to say, is propellant type is the other half of the thrust equation to pressure and nozzle throat side.. *

 

[Alby]

Anybody have a contact at Scaled Composites? Isn't SpaceShipOne powered
by an HTPB w/nitrous oxide motor? Although, it does start from 43,000/ft. :wink:

 

[r1dermon]

to my knowledge, LOX/LCH4 propellant attains a much higher ISP than that of APCP.

real world applications:
https://www.astronautix.com/props/loxlch4.htm

what do you mean by "you see why lox/methane use a lot more stuff to get the same ISP as solid APCP"

 

[ClayD]

to my knowledge, LOX/LCH4 propellant attains a much higher ISP than that of APCP.

real world applications:
https://www.astronautix.com/props/loxlch4.htm

what do you mean by "you see why lox/methane use a lot more stuff to get the same ISP as solid APCP"

Hmmmmm... I see....
While comparing apples and oragnes, I may have gotten something incorrect and/or mis-represented.

My point is why, if your going to lift something heavy, you put SRB's that use solid fuel, instead of more liquid burning engines....
the saturn iv, is the biggest antecdote to that theory..

Again, i amy be wrong in my understanding here, but propellants have an isp, then the motor has a "delivered" isp... (it takes many more pounds of liquids to equal the same newtons of force of APCP)

Sounds like a good Solid vs. liquid disucsion could be fun.

 

[ClayD]

Sounds like a good Solid vs. liquid disucsion could be fun.

So i googled it... turns out a major drawback to solid fuel is that .. it is consumed and must be replaced...

 

[jsdemar]

Just in case anyone is reading what Clay is writing, there are quite a few errors in describing Kn, burn surface, pressure, thrust curve shapes, and Isp's. Just in case.

 

[COrocket]

Hmmmmm... I see....

My point is why, if your going to lift something heavy, you put SRB's that use solid fuel, instead of more liquid burning engines....
the saturn iv, is the biggest antecdote to that theory..

Again, i amy be wrong in my understanding here, but propellants have an isp, then the motor has a "delivered" isp... (it takes many more pounds of liquids to equal the same newtons of force of APCP)

Solid motors are generally good for large amounts of thrust over short amounts of time...the shuttle SRB's give 80% of total liftoff thrust to get the vehicle going to the upper atmosphere, and the liquid engines burn much longer to get the shuttle to orbital speed...im sure there are many engineering equations that are used to figure out the best thrust over time to go into orbit, but i don't know them. Look at the Delta IV heavy...it uses additional liquid engines. It just depends what type of motor is best for the application

ISP is literally a comparison of the impulse/weight of a propellant. Sometimes engines will say theoretical ISP, which assumes perfect conditions. Since this is never the case, the delivered ISP is what actually is delivered. Then there are things like vacuum ISP. Rocket motors have an increase in thrust, for the same propellant being expelled, as ambient pressure drops as the rocket gains altitude, therfore better ISP. Usually, liquid engines will match or exceed solid propellant in ISP, depending on what fuel/oxidizer it uses. For example, don't be fooled by the fact that the SRB's are much smaller and produce much more average thrust than the liquid/main tank assembly. This does not mean they are more efficient than the liquid engines. For one, they burn much longer (need a big tank), and two, the propellant has a different density. Density has nothing to do with ISP. The SSME's, pound for pound (propellant wise) are more efficient than the SRB's, with an ISP around 400.

 

[ClayD]

Just in case anyone is reading what Clay is writing, there are quite a few errors in describing Kn, burn surface, pressure, thrust curve shapes, and Isp's. Just in case.

John, everyone has me on ignore... - But you and David Reese...

 

[cjl]

Maybe this is the blind leading the blind here, but I thought it worked a little differently. The thrust is dependent on the pressure and the throat area, and I think it's mostly independent of the propellant type.

This is pretty close to exactly right. The thrust of a motor is determined entirely by the exhaust velocity, pressure, and mass flow rate. The exhaust velocity is determined by the exit mach number and the speed of sound in the exhaust products.

The exit mach is nearly entirely determined by the nozzle - for a given expansion ratio, the exit mach will be close to the same regardless of chamber pressure, propellant type, or any other factor. The minor correction to this is that it is dependent on the ratio of specific heats of the exhaust products, but this is not going to vary very much for our propellants (or for most propellants for that matter).

The speed of sound in the exhaust products is dependent weakly on the specific heat ratio as well, but this is not the main factor. The main things that drive the speed of sound are the mean molecular weight of the gas and the temperature of the gas. The lighter and hotter the gas, the higher the speed of sound in the exhaust products (and thus the higher the speed for a given mach number). This is one reason why hydrogen fueled liquid rockets do so well - a fuel rich H2/O2 mixture has a very light mean molecular weight coupled with a fairly high flame temperature. This is the only part of the equation that does indeed depend substantially on propellant type. For the same chamber temperature and pressure, and the same nozzle, a rocket using H2 as the propellant (mean molecular weight = 2) will have triple the exhaust velocity of one using H2O (mean molecular weight = 18) (since it actually scales as the square root of the mass ratio).

The exit pressure is determined by the nozzle and the chamber pressure. For a given nozzle, the ratio of exit pressure to chamber pressure will be pretty much the same regardless of chamber pressure or propellant composition, so this is a relatively simple term once you have determined the chamber pressure of the motor. Obviously, you don't want this to be below ambient pressure (overexpanded), and ideally, you want this to be at ambient pressure (optimally expanded). This is also why a high chamber pressure is helpful - a high chamber pressure won't actually increase exhaust velocity, but it will enable you to use a nozzle with a higher expansion ratio without having the exit pressure drop below ambient. The higher expansion ratio nozzle is actually what enables high chamber pressure motors to be more efficient.

Finally, the mass flow rate is pretty much determined by nozzle throat diameter and chamber pressure. For a given chamber pressure, mass flow rate will be directly proportional to throat area (and to keep the chamber pressure constant, you want the burn surface area to throat area ratio to be about constant as well, also known as Kn). In addition, it is also determined by the gas properties - a lower density gas will have a lower mass flow rate for a given throat diameter and chamber pressure (for obvious reasons). Interestingly enough, this density dependence basically cancels with the density dependence in the speed of sound relation above (if I remember right - someone correct me if I'm wrong), so that for a given chamber pressure and nozzle, the thrust will be basically the same regardless of propellant (even through a very low molecular weight propellant will perform more efficiently, with a higher exhaust velocity and a lower mass flow rate).



(Note: this entire post is going off of memory, and it's entirely possible that there are errors, so if anyone knowledgeable like David or Bob wants to correct me on anything, feel free...)

 

[bobkrech]

In the simplest of terms, Thrust (a force) equals the propellant mass flow rate, m-dot, multiplied by the equivalent exhaust velocity.

The relatiionships between thrust, propellant mass flow rate, exhaust velocity and specific impulse are explained here. https://www.grc.nasa.gov/WWW/k-12/airplane/specimp.html

Please note that m-dot also equals Kn multiplied by the nozzle throat area multiplied by the linear propellant burn rate multiplied by the propellant density.

Bob

 

[cjl]

So i googled it... turns out a major drawback to solid fuel is that .. it is consumed and must be replaced...

As opposed to liquid fuel, which you can collect in buckets placed underneath a rocket's flight path for reuse...

(Wait a second...)

:duck:

 

[ClayD]

Thrust (a force) equals the propellant mass flow rate, m-dot, multiplied by the equivalent exhaust velocity.

Please note that m-dot also equals Kn multiplied by the nozzle throat area multiplied by the linear propellant burn rate multiplied by the propellant density.
Bob

Now... i am i reading this right...

soo, excuse my dunce or density... But does that mean the propellant density has to do with the propellant type... means the propellant type has to do with thrust generated?

 

[ClayD]

As opposed to liquid fuel, which you can collect in buckets placed underneath a rocket's flight path for reuse...

(Wait a second...)

:duck:

https://homepages.sover.net/~sbjohn/chemistry/chemx/solidvsliquid.html

:caffeine:

I have been making my own propellant for a short time... I am diligant in learning this, but some of it is plain hard.

 

[cjl]

That page isn't really a good source. Solid fuel is actually much better for long-term storage than liquid fuels are, so storability is an advantage for solids. In addition, liquid fuel boosters are pretty much never reused, while solid fuel boosters are reused already (for example the shuttle SRBs). In addition, not all liquids are throttleable, so they aren't necessarily variable thrust.

Basically, that table is very questionable, and several of its statements are just wrong. That isn't too surprising either, since it just appears to be someone's personal webpage. It does have links to sources, but its sources are broken, so they aren't helpful at all.