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Preliminary OHKLA design

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5:15 am
February 23, 2010


rpulkrabek

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I have now quickly updated OHKLA so that the body diameter is now 200 mm. I have also changed some of the lengths. I have shortened both the oxidizer tank and the fuel grain tank. This left me more room in the nose for whatever we want. I had to adjust quite a few things to accommodate these changes. For the nozzle, I basically just scaled it. I will have to run some CFD tests to verify if it still works properly or to determine the best geometry.

The mass of OHKLA (without fuel) is now 60.08 kg

The materials used were:

Nose=ABS Plastic

Nose Stopper=ABS

Connectors=6061 Aluminum

Oxidizer Tank=6061 Al

Grain Shell=6061 Al

Fuel Grain=PET (Although, it's not used for the mass calculation)

Nuts and Bolts=Steel

Nozzle=6061 Al

Fins=ABS

Here are some pictures of what it now looks like.

Ohkla reference 2010 02 23mouse

Ohkla unexplode 2010 02 23mouse

Ohkla xsection 2010 02 23mouse

Ohkla dim 2010 02 23mouse

5:22 am
February 23, 2010


rpulkrabek

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Luke Maurits said:

Does ProE have material values for cobalt by any chance?  This would be a more likely choice of nozzle material than aluminium, I think (it has a significantly higher melting point) and is considerably more dense.  The mass of the nozzle is probably a small contribution to the overall rocket mass, but we shoudl still try to get things approximately correct if we can.

EDIT: Graphite would be another good choice as well.


The current system I am using doesn't have cobalt. It's quite easy to add, though. I just need the material properties. The best place to go to is Matweb. Matweb is basically the Wikipedia for material properties. It's user generated, so you have to be cautious with the data. Here is what they have for pure cobalt. I can enter these values into ProE and then use this for the OHKLA model to determine its mass. I don't want to put this data into my work's production system, though. I will be able to enter it another place at a later time.

6:01 am
February 23, 2010


Luke Maurits

Adelaide, Australia

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Fantastic work, Ryan!  60 kg is an excellent ballpark empty mass figure, quite in line with what one might expect based on other suborbital rockets.  The new renderings look great, too.

I am curious – I know you said that the PET setting for the fuel was not used to calculate the empty mass, but can you tell me how massive the fuel grain would be with its current dimensions and port geometry, if it were made out of PET?  And also if it were made of something with a density of about 0.8 g/cm^3?  This could help us judge if we have roughly the right combustion chamber to oxidiser tank size ratio.

Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:52 am
February 23, 2010


rpulkrabek

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The current mass of the fuel grain made from PET, which has a density of 1.42 g/cm^3, is 47.957 kg.

The volume of the current fuel grain is set to 3.3772806*10^7 mm^3. If you want the mass for this from a density of 0.8g/cm^3 (which is the same as 0.8*10^-6 kg/mm^3), then the mass would be 27.018 kg

3.3772806*10^7mm^3 * 0.8*10^-6kg/mm^3 = 27.018kg

8:15 am
February 23, 2010


Luke Maurits

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Great!  This is more than enough:

With an empty mass of 60 kg, a desired delta-v of around 1500 m/s, and a specific impulse of 248 (for paraffin and N2O), the required total mass of propellant is 111 kg.  The optimal O:F ratio for paraffin/N2O is 8:1, so that means that 12 kg of that 111 kg has to be fuel (paraffin).  So 27 kg is more than twice what we need!

Of course, that 12 kg figure is an absolute bare minimum, assuming everything goes fine, and relying an an impulsive delta-v approximation.  To make sure we really get over the Karman line we would probably want to go for at least 20 kg.  We may have to make the chamber a little smaller than it appears currently, or we could just leave it as is, we'll have to wait until we have more details, but the important thing is that it seems quite certain we will not have to make it larger.  This is good, because making it larger would mean our empty mass would increase from 60 kg.

The above calculations call for a bare minimum of 111 – 12 = 99 kg of N2O.  It would be great to figure out how much internal volume we need our oxidiser tank to have to hold this much so we can check if this is on the right track, too.  I'm not sure how to go about this.  Do we store the N2O in liquid form in the tank and have it vapourise in a pre-combustion chamber, or do we store it as a high pressure gas?  If we go with the liquid route, I think the density should be fairly fixed, since liquids tend to be incompressible, but if we go with the gaseous route, the density will probably depend significantly upon pressure and temperature, so some more research would be required.  There are some figures floating around this thread along these lines, it's just a matter of following them up.

Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

8:27 am
February 23, 2010


brmj

Rochester, New York, United States

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I'm pretty sure the standard is to use liquid nitrous oxide rather than gasious. At 184.2 kelven and .1013 MPa, Wolfram Alpha says it will have a volume of about .08036 m^3. Using a diameter of 200 mm, that gives a tank 2.558 meters long. In practice, it will need to be a little longer because it won't be quite as wide as the outer diameter of the rocket, but the difference won't be huge.

Main work groups: Propulsion (booster), Spacecraft Engineering, Computer Systems, Navigation and Guidance (software)

9:39 am
February 23, 2010


Rocket-To-The-Moon

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More excellent work! It looks like a proper sounding rocket now. Can we make any estimates about the performance based on the current configuration? How much of a performance margin do we want to have to ensure that we pass 100km?

Main Workgroups: Propulsion & Spacecraft Engineering

6:42 pm
February 23, 2010


Luke Maurits

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brmj said:I'm pretty sure the standard is to use liquid nitrous oxide rather than gasious. At 184.2 kelven and .1013 MPa, Wolfram Alpha says it will have a volume of about .08036 m^3. Using a diameter of 200 mm, that gives a tank 2.558 meters long. In practice, it will need to be a little longer because it won't be quite as wide as the outer diameter of the rocket, but the difference won't be huge.


Hmm.  I am not at all certain, but I suspect that this may actually not be the norm.  184.2 Kelvin is quite cold.  At that sort of temperature one would probably need an insulated tank, cryogenic-friendly valves, etc.  All the stuff you would need for LOX, basically.  And if that were so, why would anybody use N2O instead of LOX, when LOX offers better Isp?  Also, I recall that SpaceShipOne's engine (which used N2O) had a composite oxidiser tank (there are photos of it being made on Scaled Composite's website somewhere), which suggests that they weren't doing anything at particularly low temperatures.  After some brief Googling, this page lists as an advantage of N2O over LOX: "Unlike liquid oxygen, not cryogenic. Oxidizer at -60 C or -40 C is much easier to manage. Composite tanks can potentially be used" (as an aside, the same website has another page just on N2O which looks like a good source of knowledge/safety tips, and a page on paraffin and PE wax fuels).  All of this seems to me to suggest that gasous N2O is the norm.

Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:07 pm
February 23, 2010


brmj

Rochester, New York, United States

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I just copied and pasted the values wolfram alpha gave me for liquid N2O without really looking at them, unfortunately. Space ship one used liquid N2O, supporting the idea I had gotten that that was the standard. I don't know if it is as workable in a smaller sounding rocket, though.

Now telling Wolfram Alpha to use it at 20 C, it gives me a tank about 4 meters long and 5.144 MPa.

Main work groups: Propulsion (booster), Spacecraft Engineering, Computer Systems, Navigation and Guidance (software)

11:08 pm
February 23, 2010


Luke Maurits

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brmj said:

Space ship one used liquid N2O, supporting the idea I had gotten that that was the standard.


Ah, fair enough!  I'm happy to accept that as fairly definitive.  I don't really understand well enough the interactions between pressure, temperature, density and solid vs liquid phase to think about these things clearly, I haven't had to think about it since undergrad physics.  I should probably brush up on it!

So, 4 metres long, at least very approximately.  On the current diagrams the tank looks quite a bit shorter than that.  Still, this isn't too big a problem: we've seen that we could afford to decrease the combustion chamber a little bit, and the currently allotted empty space for parachutes looks to me (admittedly based only on intuition) too large.  We should be able to find room for a roughly 4 m tank if we need it without increasing the height of the rocket much.  6 m should definitely be a workable upper limit.

I'm starting to feel like we are quite close to having the approximate dimensions and masses of most things sorted out, at least to within ballpark estimations.  The details will change, but at least we can be sure nothing is going to drastically change in size now (except maybe the fins).  The height and width of the rocket relative to the person standing next to it in the renders conveys the right impression.

This is motivating me to do more work on the pages to support propellant choice, so we can move on to more detailed considerations!  I need to focus on the SpaceUp presentation I'm supposed to be preparing for CLLARE first, though.

For what it's worth, right now I think quite strongly that N2O/Paraffin is the correct choice for our needs.  N2O because it is self-pressurising, removing the need for a Helium pressurant, and because it can be stored at sensible temperatures, removing the need for cryogen-friendly valves etc. (and overcoming what, based on readings in my hybrid rocket textbook, was a common problem for the AMROC people: LOX valves freezing stuck in half-open positions shortly after launch, leading to crashes).  Paraffin because, besides being very cheap, safe, etc., it has a very high regression rate compared to other solid fuels like HTBP (3-5 times higher!), which means that simple, single-port grain geometries can provide adequate thrust.  The "slumping" problem does worry me, but I'm sure we can come up with a solution.  The other benefits seem to make this small challenge worthwhile.

Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

1:41 am
February 24, 2010


rpulkrabek

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Here is what it looks like when the oxidizer tank is 4m long and the nose is shortened a bit.

Ohkla 4m tank dim 2010 02 24mouse

Ohkla 4m tank 2010 02 24mouse

In my opinion, this seems a bit long. What are the thoughts if we were to make the diameter 300 mm instead of 200mm?

1:53 am
February 24, 2010


Rocket-To-The-Moon

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Do you have the ability to simulate the stresses on the rocket body with FEA? If we can analyze the stresses on the rocket body and the kind of bending motions (vibration frequency, harmonic damping, ect.) then we can optimize the size to reduce the structural stress. Increasing the diameter of the rocket will also increase the propellant mass fraction without increasing the drag too much.

Main Workgroups: Propulsion & Spacecraft Engineering

2:06 am
February 24, 2010


brmj

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That helps a lot on length. It gives a tank just about 1.78 meters long.

One problem with this is that hybrid rocket fuel ports like to be rather long compared to their width, makeing shorter single-port designs problematic. I'm not sure quite what the limits are for that, though, and a multiport geometry would render this irrelivant.

Main work groups: Propulsion (booster), Spacecraft Engineering, Computer Systems, Navigation and Guidance (software)

2:13 am
February 24, 2010


Rocket-To-The-Moon

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Post edited 2:14 am – February 24, 2010 by Rocket-To-The-Moon


I remember someone mentioning the optimum length to diameter ratio. 7:1 stands out in my head (for a single port configuration), but I can't find the exact place this was mentioned.

Here are the references to lenght vs diameter.

Main Workgroups: Propulsion & Spacecraft Engineering

2:47 am
February 24, 2010


DenisG

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I'm currently reading a book about rocketry and the width-to-length ratio is one of the subjects covered pretty well. I can't tell anything definite right now though, since I'm in a hurry. I'm writing an exam tomorrow, but after that I can report back on it with math. BTW, the nose cone of ABS still seems unrealistic to me. I believe it should be made at least of aluminium because of temperature and pressure; I'll look into that in detail as well.

3:47 am
February 24, 2010


rpulkrabek

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Rocket-To-The-Moon said:Do you have the ability to simulate the stresses on the rocket body with FEA? If we can analyze the stresses on the rocket body and the kind of bending motions (vibration frequency, harmonic damping, ect.) then we can optimize the size to reduce the structural stress. Increasing the diameter of the rocket will also increase the propellant mass fraction without increasing the drag too much.


Yes, this is very easy to do. We just need to know what stresses/forces and constraints will be involved. 

3:50 am
February 24, 2010


rpulkrabek

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DenisG said:

I believe it should be made at least of aluminium because of temperature and pressure; I'll look into that in detail as well.


The material can easily be changed. I think after SpcaceUp we should put more considerations towards our design decision process. We have done basic decisions, but every that has been decided has been only to provide a concept or something that looks presentable. 

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