expansion ratio's...and other fun stuff

[butalane]

ClayD,

Your posts are riddled with poorly drawn conclusions and false information. If you're going to reach your own made up conclusions and put words in other people's mouth, it is unclear why you are bothering to even engage in this discussion. I only post this because the OP was looking for valid information, and I hope he understands that the conclusions you draw, statements of fact you make and heresey you spread are virtually all fallacious.

To add something to the discussion:

Erosive burning is something that exists in all rocket motors with high cross flow. Even the SRB experiences this (https://ntrs.nasa.gov/search.jsp?R=19830026740&qs=Ns=Loaded-Date|1&N=4294655563) The degree to which the phenomenon exists varies based on configuration and propellant formula, and sometimes can be ignored. The results of erosive burning do not only show themselves in the form of grain stripping but also in their effect on burning rate and subsequently thrust curves. Turbulent mixing and reduction in the distance of the flame front from the propellant burning surface causes an increased burn rate in grains with higher mass flux than others. This can disrupt what would be a normal burning rate curve by cause a spike at ignition and a premature dip near burn out as lower grains burn out before the upper grains.

A little more on erosive burning:
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20030066237_2003075468.pdf

Nozzleless motors do experience a higher degree of erosion than motors with nozzles. Read:
https://serge77.rocketworkshop.net/nozzleless0/nozzleless.pdf

This might be helpful as well:
https://en.wikipedia.org/wiki/Choked_flow

 

[ClayD]

virtually

In reality, I give little weight to your opinion of me, and and less to your justification or judgement on my activites.

Perhaps you are more elequant, and much more intelligably derived in your explanation than I was, and you have an exceptional presentation. sofa king what. I am not a "scholarly idevidual", wasnt making a thesis statement or professing and alpha-omega position of something i am learning. Nor was I writing a book.

If someone takes a fecal application and false pretence from the internet, and applies it in hopes of thinking its correct - Shame on them....

 

[r1dermon]

thanks for your post.

just a lay-mans question here...you say erosive burning is a product of high cross-flow. this makes sense, however, let me pose a question.

in a rocket motor, not taking into account a specific propellent, with a single, solid grain and a full core, would that motor be pretty much unsucceptable to erosive burning characteristics since it doesn't exhibit any cross flow (in between grains), assuming it burns normally (ie, doesn't burn through the liner, or separate the grain from the case, assuming case-bonded).

in other words, is erosive burning a characteristic only found in motors containing more than a single grain of propellent?

also, clayD, please try and keep this thread on topic. my goal is to log a bunch of information about the physics of why and how solid rocket propulsion systems work, and what is done to optimize performance through changing characteristics of the geometry of the design. whether it be the dynamics of the nozzle design, grain design, case design...etc...the chemistry of why they work is not what im looking for. more specifically, im wondering how (in my head) to produce the most efficient propulsion system using x-propellent formula. for instance, to create a rocket motor that harnesses every bit of energy as possible, which is stored within the propellent. does that make sense?

the reason i started with expansion ratio's is because it's very interesting to me, that so much performance can be gained or lossed, based on the design of the nozzle.

 

[ClayD]

thanks for your post.

just a lay-mans question here...you say erosive burning is a product of high cross-flow. this makes sense, however, let me pose a question.

in a rocket motor, not taking into account a specific propellent, with a single, solid grain and a full core, would that motor be pretty much unsucceptable to erosive burning characteristics since it doesn't exhibit any cross flow (in between grains), assuming it burns normally (ie, doesn't burn through the liner, or separate the grain from the case, assuming case-bonded).

in other words, is erosive burning a characteristic only found in motors containing more than a single grain of propellent?

also, clayD, please try and keep this thread on topic. my goal is to log a bunch of information about the physics of why and how solid rocket propulsion systems work, and what is done to optimize performance through changing characteristics of the geometry of the design. whether it be the dynamics of the nozzle design, grain design, case design...etc...the chemistry of why they work is not what im looking for. more specifically, im wondering how (in my head) to produce the most efficient propulsion system using x-propellent formula. for instance, to create a rocket motor that harnesses every bit of energy as possible, which is stored within the propellent. does that make sense?

the reason i started with expansion ratio's is because it's very interesting to me, that so much performance can be gained or lossed, based on the design of the nozzle.

Cross flow, is down the core, not end burning.(across the burning surface)
a long grain, generally is more susceptable to erosive burning. Short "bates" geometry, is somewhat more turbulent, and you get less cross flow.
(ie, more neutrual grain, diameter to lenght ratio.)
i am sure some will take offence an idiot presented this in a fecal manner...

erosivity, --- erosion -- like rivers errode channels in the earth..

the way i look at erosivity, is that particles are entering the cross flow, not fully burned. they prematurely break away from the grain.
essentialy increasing the total burning surface of propellant.

 

[r1dermon]

so cross flow is the flow specifically down the core?

the way i understood it, cross flow is the flow between the grains which essentially flow horizontally into the core. so this was a misconception?

 

[ClayD]

so cross flow is the flow specifically down the core?

the way i understood it, cross flow is the flow between the grains which essentially flow horizontally into the core. so this was a misconception?

no, cross flow is not down the core....

Cross flow is across the burning surface.

In a long single grain motor, cross flow just happens to be down the core.

Port flow is down the core.

So between the ends, the flow going into the the core, is cross flow of the end of the grain.

thats my understanding of it... however wrong.

https://www.jamesyawn.net/clearpvc/index.html
some neat stuff that is fairly new (i dont remember seeing it... go watch some videos..

 

[r1dermon]

so cross flow is gaseous flow perpendicular to the propellent...thats what i gather.

 

[jsdemar]

the core id was .5" the legnth was 21.76.
...well onsite the 1 made me manufacture 1.3 grams, and i essentially discarded the more than half...

1/4 gram is enough. And not packed in anything rigid. Just plastic wrap. About twice as long as the core.
https://thrustgear.com/data/thermite_mass.JPG

 

[ClayD]

ClayD,

Your posts are riddled with poorly drawn conclusions and false information. If you're going to reach your own made up conclusions and put words in other people's mouth,


Erosive burning is something that exists in all rocket motors with high cross flow.

Turbulent mixing and reduction in the distance of the flame front from the propellant burning surface causes an increased burn rate

the burn is presumed to be perpendicular to the burning surface, if there is cross flow, this presses down the flame front onto the surface increasing the temperatures. the flow is across the burning surface, cross flow for short.

the higher velocity the flow, the more the cross flow (higher cross flow) is to the perpendicular burning front.

End burners are the best example of nearly no cross flow. the mass flux is flowign away from the grain.

Cross-flow, is used to determine erosive enhancement in "simulations".

https://etd.gatech.edu/theses/avail.../unrestricted/mcdonald_brian_a_200407_phd.pdf

 

[jsdemar]

so cross flow is gaseous flow perpendicular to the propellent...thats what i gather.

Yes, flow across the face of the propellant core. As the gas flow travels down the core, it increases in density and velocity. The measure of "mass flux" combines these based on the port area. The worse spot for mass flux is right at the bottom end of the bottom-most grain.

The situation is worse then there isn't choked flow, such as a nozzleless motor or a core area <= nozzle throat area.

Erosive burning is worse under acceratation. It's also worse depending on solids loading and particle sizes.

 

[ClayD]

so now that i have read a little bit on break, i will issue said retraciton:
it has nothing to do with erosion.. like my soo sadly drawn conclusion...

5.6 Thermal Boundary Layer Profile Dependency
The primary mechanism effecting erosive burning is increased heat transfer from the free-stream, through the boundary layer, down to the propellant surface.

5.7.0 Finite Rate Chemistry Model Overview
The primary mechanism in erosive burning is the augmentation of transport
properties in the boundary layer due to turbulent mixing.

 

[cjl]

The situation is worse then there isn't choked flow, such as a nozzleless motor or a core area <= nozzle throat area.

There is still choked flow happening in those cases, but the choke point is the aft end of the rear grain, rather than the nozzle throat (which explains the high erosivity).

 

[jsdemar]

There is still choked flow happening in those cases, but the choke point is the aft end of the rear grain, rather than the nozzle throat (which explains the high erosivity).

I guess it's a point of definition... restrictive choked flow vs straight choked flow. With classic choked flow, it reaches equilibrium conditions faster and under stable conditions. A straight pipe (no nozzle) will hit sonic flow at the opening of the pipe under certain conditions (long enough, adiabatic, compressible). But the equilibrium condition is unstable, sending shock waves back into the pipe. If there is incomplete combustion (erosivity makes that more likely) it may choke before the opening, or go supersonic within the pipe. Nasty conditions.

 

[r1dermon]

i was recently informed of a rocket motor (L950) which had a smaller core diameter than nozzle throat.

the youtube video of it functioning can be found here:
https://www.youtube.com/watch?v=H5T8o1VxYBU&feature=player_embedded

upon an inquiry, the builder stated that the core was smaller in diameter than the nozzle throat, and yet the motor functions fine...

so my question is, is this predictable behavior? did this motor function despite the fact that the core diameter was smaller than the nozzle throat?

even if efficiency is lost initially due to the design, could the added total impulse be a net positive despite the negative of the poor efficiency during the initial burn of the motor?

is this just blind luck that this motor functioned? or would there be a net positive in some applications where the core is smaller than the nozzle throat? in other words, despite not being preferable most of the time, is there any time that it would be a better option?

thanks.

 

[ClayD]

R1der, this guy can give you hand on throat / port
https://www.tdkpropulsion.com/2010/04/port-to-throat/

 

[daveyfire]

is this just blind luck that this motor functioned?

Nope, it's engineering. (I'd post the XKCD science comic again but bobkrech might get mad )

or would there be a net positive in some applications where the core is smaller than the nozzle throat?

Absolutely.

in other words, despite not being preferable most of the time, is there any time that it would be a better option?

I've (intentionally) done two 98/17,500 loads with port:throat area ratios at or less than 1. I know Fred Azinger does it (also in the 17.5, which is kinda long) on a pretty regular basis. Others probably, too. (e.g., the N10,000 is awful close.) With consistent propellant and a tested design, there's no reason for it not to work.

Such a design can be of special advantage in case bonded/glued motors, so you can get the mandrel through the nozzle and still have room to pour the propellant in.

 

[cjl]

Others probably, too. (e.g., the N10,000 is awful close.)

The N5800 is pretty close too. I'm not sure if the port is actually smaller than the throat or not, but it's definitely close.

 

[bobkrech]

R1der, this guy can give you hand on throat / port
https://www.tdkpropulsion.com/2010/04/port-to-throat/

Excellent reference.

 

[r1dermon]

Thanks clayd...echoing jobs sentiment, that's exactly what i was looking for.

Thanks for all the responses as well.

 

[r1dermon]

The flame looks like Wayside White. The following is speculation based on that guess.

Recall Ryan's post on pressure loss down the port from mass addition. Combine that with extremely high port velocities from your grain design. Horizontal mass flow injection from the BATES grain segment faces increases port velocity further, as the vena contracta is further reduced in size by this flow. The result is a port velocity that is incredibly fast, and thus a port static pressure that is incredibly low compared to stagnation conditions at the head end and near the case wall.

With soft propellants (like WW, which is a low solids loading), such a pressure differential can be enough to "suck" the propellant walls inward slightly. This reduces the flow area further, increasing the port velocity, and reducing the static pressure. The result is an unstable feedback loop which will end when the propellant grain rips itself to -- in this case -- "dime size" shards. These appear to have then been ejected from the nozzle.

So that's what I think happened. The Ap/At rule of thumb is just that -- a rule of thumb -- but until you (or whomever designs your motors) take full control over your design and have an understanding of all the reasons the rule of thumb is what it is, it's best not to mess with it.

So you're saying that head end pressure is higher than aft end? If yes, is this expressed in a law of physics?

 

[cjl]

Yes, the head end pressure is higher than the aft end. This is because there is flow from the head end to the aft end, and there is no such thing as a flow without losses. For the flow to travel, there must be a pressure gradient. How much of a pressure gradient can vary widely, but the pressure at the head end must always be slightly higher than the pressure at the nozzle, or else there would be no flow.

 

[daveyfire]

So you're saying that head end pressure is higher than aft end?

Yes, depending on the flow Mach number at the aft end of the grain. See, e.g., the data set attached from a friend's thesis. The original was in color, but this black and white version still communicates the pressure difference between aft and head end for a port Mach number, in this case, of 0.765.

If yes, is this expressed in a law of physics?

It's explained by gas dynamics -- check the link to my blog that ClayD posted previously. (Which I'm extremely happy, btw, was able to pass Bob's muster -- I'm so glad I didn't make any more egregious typos! )

 

[Chuck Rogers]

Yes, the head end pressure is higher than the aft end. This is because there is flow from the head end to the aft end, and there is no such thing as a flow without losses. For the flow to travel, there must be a pressure gradient. How much of a pressure gradient can vary widely, but the pressure at the head end must always be slightly higher than the pressure at the nozzle, or else there would be no flow.

See pages 50-54 in the Departures from Ideal Performance Technical Report Download on the RASAero web site for the theory and equations. The flow loss is the total pressure loss between the head end of the core and the aft end of the core and the nozzle entrance. While of interest for understanding what is happening in the core of the solid rocket motor, it is also an important effect for correcting chamber pressure data (measured at the head end of the motor) when what you really want is the nozzle stagnation pressure, which is at the aft end of the motor. Figure 13 on page 50 really illustrates the concepts of what is being discussed here. And it is tied to the ratio of the port cross-sectional area to the throat cross-sectional area, see Figure 16 on page 53.


And for the erosive burning subject in general, see the Erosive Burning Design Criteria Technical Report Download on the RASAero web site in the Solid Rocket Motor section.


Chuck Rogers
Rogers Aeroscience

 

[r1dermon]

before i write anything, i just want to say, this (for me) has been the most informative thread i've ever read on this website. thanks to all who have participated.

now back to your regularly scheduled programming...

obviously, what people read on the internet doesnt make it "true"...and one of the reasons im asking the question about the pressure difference between the head end, and aft end of a motor, is because of this guy...

https://www.scss.tcd.ie/Stephen.Farrell/ipn/background/Braeunig/propuls.htm

The pressure distribution within the chamber is asymmetric; that is, inside the chamber the pressure varies little, but near the nozzle it decreases somewhat.

so what you guys are telling me is, this information is either incorrect, or mildly inaccurate? my understanding is, although the pressure may vary "little", it's all relative really, and little doesn't mean "not at all". any insight?

 

[daveyfire]

so what you guys are telling me is, this information is either incorrect, or mildly inaccurate? my understanding is, although the pressure may vary "little", it's all relative really, and little doesn't mean "not at all". any insight?

"Overly simplified" is probably the best way to put it. In an ideal rocket motor, nozzle static pressure is head end static pressure is stagnation pressure is chamber pressure -- but this isn't an ideal world, unfortunately. "little" can be quantified a priori using the equations in Chuck's document or in my post.

While of interest for understanding what is happening in the core of the solid rocket motor, it is also an important effect for correcting chamber pressure data (measured at the head end of the motor) when what you really want is the nozzle stagnation pressure, which is at the aft end of the motor.

This is a really good point that I've never thought about previously, but definitely should. You can get hosed when calculating thrust coefficients and C* efficiencies in long motors when using uncorrected head-end data.

Chuck, I have to say - I read the original erosive burning series as a high school senior back when it first came out in HPR (2005), and I've read it again every year since, understanding a little bit more each time as my knowledge improved. Finally, reading back over it now after a bunch of time thinking about it, it makes sense, and I've come to realize is an extremely well written and assembled summary of the topic. Thank you for putting it together!

 

[ClayD]

obviously, what people read on the internet doesnt make it "true"...and one of the reasons im asking the question about the pressure difference between the head end, and aft end of a motor, is because of this guy...

https://www.scss.tcd.ie/Stephen.Farrell/ipn/background/Braeunig/propuls.htm

David posted on the "overly simplified"...
there is also an Out-of-context issue going here. Not all rocket motors are aluminum pipe. (or long cylinders..) Some are liquid motors with more spherical physical shape.. Since were talking pressure gradiants, shape somewhat comes to play... i read some stuff a while back , that the nozzle is best when it is a function of the "centered of the area of the combustion chamber with the highest pressure."(meaning a ball with a divergant cone comming from near the center....)


Chuck, I also agree with David, thanks for puting that all together (and sharing it.)

Hey David, thanks for your "research" sharing as well...

 

[AlphaHybrids]

Liquid motors are quite different. They use L* to get the chamber volume you need to combust the reactants efficiently. You also create your flow by injecting at a higher pressure (thus creating your pressure gradient) than chamber pressure.

Edward

 

[ClayD]

Liquid motors are quite different. They use L* to get the chamber volume you need to combust the reactants efficiently. You also create your flow by injecting at a higher pressure (thus creating your pressure gradient) than chamber pressure.

Edward

"Think Mcfly" comes to mind as I chuckle...

It would indeed stand, that your injector oraface has to be a higher pressure than your chamber pressure.. and that it would generate its own pressure gradiant...And... most liquid motors have the injector on the "head" end.

 

[Chuck Rogers]

obviously, what people read on the internet doesnt make it "true"...and one of the reasons im asking the question about the pressure difference between the head end, and aft end of a motor, is because of this guy...

https://www.scss.tcd.ie/Stephen.Farrell/ipn/background/Braeunig/propuls.htm



so what you guys are telling me is, this information is either incorrect, or mildly inaccurate? my understanding is, although the pressure may vary "little", it's all relative really, and little doesn't mean "not at all". any insight?

The quote from the web page is:

The pressure distribution within the chamber is asymmetric; that is, inside the chamber the pressure varies little, but near the nozzle it decreases somewhat.

The web page is written for a generic rocket engine/motor, but I believe what he is really referring to is a liquid rocket engine.

Figure 17 on Page 54 of the Departures from Ideal Performance Technical Report Download on the RASAero web site shows the nozzle stagnation (total) pressure divided by the pressure measured at the injector face for a liquid rocket engine (the original source of the graph was NASA SP-125) as a function of the ratio of the combustion chamber cross-sectional area to the nozzle throat area. Unless you have a nozzleless or nearly nozzleless liquid rocket engine combustion chamber, for realistic combustion chamber cross-sectional area to nozzle throat area ratios the loss in total pressure from the injector face to the nozzle stagnation pressure is small (although it is not zero, and should be corrected for to turn the pressure measured at the injector face into the chamber pressure for the rocket engine).

Note though that this is for a liquid rocket engine, not the flow down the core into the nozzle of a solid rocket motor. The web page author probably should have noted that what he was referring to was a liquid rocket engine.


Chuck Rogers
Rogers Aeroscience

 

[Chuck Rogers]

Chuck, I have to say - I read the original erosive burning series as a high school senior back when it first came out in HPR (2005), and I've read it again every year since, understanding a little bit more each time as my knowledge improved. Finally, reading back over it now after a bunch of time thinking about it, it makes sense, and I've come to realize is an extremely well written and assembled summary of the topic. Thank you for putting it together!

You're very welcome. Those Tech Articles were written to be read, as they say, with the information within them to be put to good use.


Chuck Rogers