Forum | Preliminary OHKLA design

What are the physical dimensions of this rocket? I see that you have scales on the images, but it is hard to tell accurately. 143 kN is a lot of thrust (as much as an F-16 fighter jet)!

Have you spoken with anyone regarding fabrication of parts and pieces? If you can get a local machinist interested you could probably expect him to help mill out the parts.

Post edited 12:43 am – February 9, 2010 by rpulkrabek

Rocket-To-The-Moon said:

What are the physical dimensions of this rocket? I see that you have scales on the images, but it is hard to tell accurately.

Have you spoken with anyone regarding fabrication of parts and pieces? If you can get a local machinist interested you could probably expect him to help mill out the parts.

The scale from the pictures isn't really meant to accurately measure the rocket as much as it is to give an idea of what the size is. I could create nice 2D CAD documents with the measurements shown, but at the moment, I only have my work's default template with their logo and such. I'll see if I can come up with a modified one. Maybe at some point, we can have an official CSTART CAD template. This will have to wait until a later time when things are more complete.

For now, I can provide screenshots with quick measurements. The measurements of the rocket as a whole should not be taken seriously. They are just guesses based on what others have done. We should take the approach Luke had mentioned and use the equations to determine the radii and lengths needed to provide the delta-v we desire. All measurements are in mm.

The nozzle geometry is what I spent my time working on and has more accurate dimensions. Keep in mind, however, the dimensions will change based on what the dimensions will be for the rocket itself and the combustion pressure that occurs.

To answer your second question, no, I haven't asked anyone about machining these parts. This would have to wait until we are ready to start testing, I believe. Perhaps brmj could even do some of this from his university. A static test would be a great start once we know the fuel grain geometry and fuel composition we are using. That sounds quite exciting; I wonder how far away we are from that.

Post edited 1:42 am – February 9, 2010 by brmj

I can pretty much guarantee that the thrust numbers are bogus in the context of what OHKLA will actually be getting. This image shows a thrust graph from a HATV test. They get nothing near 143 kN. Something is obviously going wrong, but the fact that the flow patterns we are getting line up with what is expected suggests something is also going right. Just on a hunch, would you mind rerunning it with lower chamber pressures? Typical chamber pressure seem to scale with rocket size as far as I can tell, and the upper stage engine in a Falcon 1 runs at .93 MPa. I'm not quite sure what we can reasonably expect, but 3 MPa may well be too high.

I've been wikipediaing rocket engines pretty heavilly, and these numbers just aren't lining up. I'm not sure exactly what the underlying issue is, but there seems to be one.

If I remember correctly, the pressure number was arbitrary.

I'm not sure if the issue of the flow through the grain would matter, or if the postcombustion chamber would even it all out. In this particular model, however, the grain doesn't seem to do anything interesting to the flow, so I'm not sure how that could make a difference. I might have missed something, however.

I don't realy know what I'm doing here, unfortunately, but I suspect that there is some model we could use for sanity checking. I suppose I could go ask the Project Meteor people what they do for this sort of thing. Another option might be to ask the Copenhagen Suborbitals for a bunch of HATV data and, if we get it, compare our results to their actual static tests and, if they are grossly off, try an figure out why.

3 MPa was only a guess to what chamber pressure to use. I am not quite sure what our fuel of choice will produce. I can surely run a new set of simulations for a lower combustion pressure. What pressure should I use, 1 MPa? 0.5 MPa?

As for the fluid flow through the fuel grain, I can reassign the "inlet" as well. At the moment it as at the very top of the fuel grain housing. Is it better to have the inlet at the end of the fuel grain, or would it be even better to have the inlet assigned to every surface of the fuel grain?

I don't know, something in that range I guess. Let's arbitrarily say perhaps 0.5 MPa, unless someone else has a reason to favor some particular value. Unfortunately, we don't have good numbers to estimate this off of yet as far as I can tell.

As for the fluid flow through the fuel grain, I can reassign the "inlet" as well. At the moment it as at the very top of the fuel grain housing. Is it better to have the inlet at the end of the fuel grain, or would it be even better to have the inlet assigned to every surface of the fuel grain?

The absolute most accurate way to do this would be to have some come off of the grain, and some off of the top of the housing, since there will be an excess of oxidizer in the gas mix. That said, I am of the opinion that this shouldn't be a big factor in what it does beyond the grain.

I am unable to perform the simulations today, it will have to wait until maybe tomorrow

Okay, no hurry. You've been doing a lot of great work. Thanks.

Luke Maurits said:

Okay, time for some numbers!

At this Wiki page (which will hopefully slowly grow with time to be a good repository of hybrid rocket info) I have began transcribing a table of propellant info from the hybrid rocket textbook I currently have on loan.  Please note that this is a small portion of the entire table – I have focused on the propellants that seem most likely for us because I don't have too much spare time at the moment.  I will endeavour to complete the table in time.

Great, this will help out quite a bit! My only suggestion is that we keep the units metric. Also, to avoid confusion, it might be a good idea that we start writing pressure like kPaa or kPag for absolute and gauge pressures. Thanks for starting this, I would like to think that over time, these types of tables will be quite complete.

Luke Maurits said:

To get the propellant mass figures I will need an estimate of the empty rocket mass.  Anybody want to throw out some figures?

Do you need the mass of the rocket minus the oxidizer and fuel grain? If so, with every component currently assigned as steel, the Pro/E model I have gives the following information with a coordinate system at the center of the nozzle diameter (the very bottom of the rocket):

VOLUME =  4.8727940e+08  MM^3

SURFACE AREA =  2.6528419e+07  MM^2

AVERAGE DENSITY =  7.8270800e-06 KILOGRAM / MM^3

MASS =  3.8139749e+03 KILOGRAM 

CENTER OF GRAVITY with respect to DEFAULT coordinate frame:

X   Y   Z     0.0000000e+00  2.6753948e+03  0.0000000e+00  MM

INERTIA with respect to DEFAULT coordinate frame:  (KILOGRAM * MM^2)

INERTIA TENSOR:

Ixx Ixy Ixz  4.2233764e+10  0.0000000e+00  0.0000000e+00
Iyx Iyy Iyz  0.0000000e+00  3.4322134e+08  0.0000000e+00

Izx Izy Izz  0.0000000e+00  0.0000000e+00  4.2233764e+10

INERTIA at CENTER OF GRAVITY with respect to DEFAULT coordinate frame:  (KILOGRAM * MM^2)

INERTIA TENSOR:

Ixx Ixy Ixz  1.4934334e+10  0.0000000e+00  0.0000000e+00

Iyx Iyy Iyz  0.0000000e+00  3.4322134e+08  0.0000000e+00

Izx Izy Izz  0.0000000e+00  0.0000000e+00  1.4934334e+10

PRINCIPAL MOMENTS OF INERTIA:  (KILOGRAM * MM^2)

I1  I2  I3   3.4322134e+08  1.4934334e+10  1.4934334e+10

ROTATION MATRIX from DEFAULT orientation to PRINCIPAL AXES:

0.00000        0.00000        1.00000

1.00000        0.00000        0.00000

0.00000        1.00000        0.00000

ROTATION ANGLES from DEFAULT orientation to PRINCIPAL AXES (degrees):

angles about x  y  z   0.000         90.000         90.000 

RADII OF GYRATION with respect to PRINCIPAL AXES:

R1  R2  R3 2.9998409e+02  1.9788096e+03  1.9788096e+03  MM

Luke Maurits said:

Oh, another thought on explaining the absurdly high thrust estimates: to get a thrust estimate from the gas exit velocity and the nozzle area, we naturally also need a gas density.  This will certainly depend on the choice of propellants.  What have we been using for this so far?  Due to the high temperatures involved, I expect the density will be quite low.  If the simulations so far have been using standard density of air or something, that could explain things.

I have been using N20 as a combusted gas. I am not sure what the density of it was though. If I remember, I can find out more specifics when I have access to a work station tomorrow.

I should also point out that the entire enclosure the model was simulated in was N20. For a future simulation, it would be nice to have the enclosure as air and the combusted fluid as N20. It might also be beneficial to have another inlet of air at the front of the rocket.

Post edited 6:57 am – February 9, 2010 by rpulkrabek

Luke Maurits said:

First of all, it's really awesome that your software can simply give us mass etc. figures like this.

Second of all, 3 tonnes?!?!?!  Even using steel, that sounds like way too much to me.  Have I missed something?

After reading how you said that, I was a bit shocked too. Then after thinking about it logically, yes, I guess it is about 3 tonnes of steel. As an example, I'll go through the steps of the mass for the oxidizer tank, which is roughly 3 meters tall and has an outside diameter of 255mm and an inside diameter of 215mm.

l=3000mm

r_o=255mm

r_i=215mm

d_steel=7.85g/cm^3 => d_steel=7.85*10^-6kg/mm^3

V=l*pi*r_o^2-l*pi*r_i^2=1.7719*10^8mm^3

m=V*d_steel=1390.9kg

So, the oxidizer tank alone has a mass of about 1.4 tonnes. According to Pro/E, I am given the mass to be 1.4096042*10^3kg.

The values from Pro/E are correct, it's the design of the rocket that needs rethinking. Do we go with steel as the material or do we choose something better? I am sure the wall thicknesses don't need to be as thick either.

Luke Maurits said:

As for using N2O as the gas, this may be the problem.  While some uncombusted N2O may make it out of the nozzle, much more of the exhaust gas will be composed of the byproducts of the combustion reaction.  N2O's role in the reaction is as an oxidiser, so I would suspect that straight nitrogen gas, N2, is part of what comes out the other end (and obviously N2 is less dense than N2O).  The fuel quite likely contributes something to the exhaust gas, too.  CO2 and H2O seem like likely candidates to me, but I'm not a chemist.  Even if we are able to figure out the exact mix of gases that comprise the exhaust, we won't be able to reckon an accurate density without knowing the combustion temperatures either.  Fortunately, finding combustion temperatures for common propellants shouldn't be too hard.  There may even be tables of this in the book if I look around.

Good point. I'll try to investigate further into what types of material properties can be used. I do remember seeing some sort of combustion branch in the setup tree, so it could be possible that we can specify the material involved in the combustion and the software will solve the rest, but this is also probably wishful thinking.