Question A question about fuel/oxidizer ratios for different fuel choices...

[ISProgram]

My first thread... Planning on making a HLV addon (in the long term) to go with a fictional space program.

I've actually already designed a model with the dimensions of said HLV, though I'm having some problems calculating the fuel and oxidizer tank lengths. The aforementioned HLV uses a RP-1/LOX for its first stage and LH2/LOX for its second stages(s).

Basically, I have a question regarding the fuel to oxidizer ratio for rockets. Mainly, exactly what is that ratio? I know of the fact that rockets can have vastly different tank sizes even if they are the same diameter/length because of the different densities of preferred fuels and oxidizers (LH2/LOX or RP-1/LOX).

My knowledge on spaceflight (and programming) is very basic.

Any feedback would be appreciated...

 

[boogabooga]

Assume that you are not running oxidizer rich or fuel rich.

At the most basic level, you need to know the ratio of oxidizer to propellant as it appears in the combustion chemical equation. For example:

2H2 + O2 -> 2H2O

This means that you need 2 MOLES of hydrogen for each MOLE of oxygen.

http://en.wikipedia.org/wiki/Mole_(unit)

Use the formula weight to convert from moles to mass. This will give you the ratio of the MASS of fuel and oxidizer

Then, you can use mass and known density of fuels to calculate the ratio of tank volumes.

RP-1 is tougher because it is a complex mixture rather than a pure substance, but you should be able to find an average formula weight somewhere.

 

[ISProgram]

Thanks. With regards to the tank ratio, I drew some inspiration from some rocket cutaway.

The top one burns LH2/LOX and was based off tank ratios from SLS. The bottom was based off a couple other rocket cutaways and burns RP-1/LOX.

Assume that you are not running oxidizer rich or fuel rich.

I assuming that if the rocket engine is, for example, fuel-rich, the fuel tank would have to be proportionally larger compared to the oxidizer to accommodate this extra fuel, and the reverse for oxidizer-rich.

This of course may make the rocket larger, and if so, exactly what is about being "rich" worth it? Increased efficiency?

 

[Loru]

This is post I've written for one of the orbinauts. It may be helpfull for you too.

Part1: For Solid stages. Cylinder volume/height calculations apply also to liquid stages.

1.) First you need is total volume of your fuel. For that you'll need density of your fuel. Let's assume typical solid fuel is 1.35 g/cc or 1350kg per cubic meter (m3).

So: 392kg / 1350kg/m3 = 0.3 m3 - not much but we have to remember that solid motors have a hole to expand "burning surface" and increase thrust. Let's add 60% of volume for the hole: 0.3 m3 * 1.6 = 0.48 m3 This is your target volume of fuel case.

2 Now we need to determine height. It's simple sterometrics (3d geometry)

Propelant case is a cylinder so we volume algorithm for cylinder:
V (volume) = Pi*radius^2 * height - we transform this algorithm to calculate height.

Height =Volume / Pi*(radius^2)

H= 0.48m3 / 0.28 m^-2
H= 1.72m - here is your height (nozzle not included so you have to add like 0,7 to 1 meter. )


Part 2 Liquid stages.

First you need to determine your mixture ratio - typical ratio for RP1/LOX is around 3 but remeber it's Ox/Fuel so we have 3 parts of Ox per 1 part of Fuel here. (Calculations are done for this fuel ratio - for different ratios and mixtures you have recalculate data))

LOX density is around 1g/ccm and RP 1 is 0.806 g/ccm

Give your data:

Dry mass: 910 kg
Fuel mass: 8276 kg
Diameter: 1.6 m
Fuel: LOX/RP-1

We have: 6207kg of LOX and 2069kg of RP-1. Which gives us ~6.2 m3 for LOX tank and 2.6 m3 for RP-1 tank.

As it's liquid stage with insulation let's assume 10cm of insulation so it leaves us with 1.5 meter for LOX tank (0.75m radius).

So:
LOX Height of tank: 6.2 m3 / 3.14*(0.75m^2) = 3.52 m
RP-1 Height of tank: 2.6 / 3.14*(0.8m^2) = 1.3 m

sum gives us 4.82 but let's add ~10% for tank endings etc. so let's assume 5.5 meters of tanks. Now you have to add room for equipement and engine.

I hope it's clear enough and will help you.

Cheers

Here is excell spreadsheed I'm using for rocket calculations. (just enter the data into white fields.

 

[Dantassii]

Fuel comments from a Rocket Scientist

One of the things that also determines the fuel to oxidizer ratio of any rocket engine is the temperature of the combustion. All Hydrogen/Oxygen rocket motors run either fuel rich or oxidizer rich for this reason. Burning H2/O2 at stochiometric conditions results in a temperature so high that no material could be used as a combustion chamber for the propellant flows necessary for a rocket engine. During a project I worked on during my Ph.D. attempt, our design team discovered that the closer you could get to perfect combustion the more efficient your rocket motor was, but only up to a point. Beyond that point, you had to spend so much rocket motor mass on cooling that the overall performance of the rocket decreased rapidly.

Dantassii
HUMONGOUS IMS shipbuilder
B.S. Aerospace Engineering
M.S. Mechanical Systems Engineering
3 years towards Ph.D. Aerospace Engineering

 

[ISProgram]

Ok, some data I’m just going to throw out there in case anyone’s interested. The data is that of the revised first stage tanks. Both have a diameter of 7.6 m.

LOX tank length: 17.94 m overall, cylindrical (13.06 m) and tank domes (2.44 m).

RP-1 tank length: 32.23 m overall, cylindrical (27.4 m) and tank domes (2.44 m).

As a further sense of scale, the person to the left is about 1.75m tall.

The first stage will probably have a dry mass less than that of the SLS core stage, given that it has reduced diameter and height, as well as different engines and components. Wet mass, however, is a different story, and will probably be higher, on the basis that RP-1 has a greater density than LH2 and the differently sized LOX tanks.

Any (minor) additional information is at http://danirevan.deviantart.com/art/Hermes-Common-Core-Stage-372093380


Thanks for the support so far. :thankyou:

---------- Post added 03-02-14 at 04:24 PM ---------- Previous post was 03-01-14 at 08:43 PM ----------

Made a few mistakes with the dimensions, some of which had persisted for a few MONTHS. :facepalm:

First, the diameter is 7.91 meters, not the previously claimed 7.6 meters. It's diameter is thus 25.95 feet now. This of course also applies to the diameter of the tanks as well.

I also reversed the dimensions for the LOX and RP-1 tanks. The LOX tank is larger, at 32.23 meters in length, while the RP-1 tank is 17.94 meters.

The intertank is 5.44 meters is length and 7.91 meters in diameter.


Also, preliminary calculations show that the RP-1 tank has a capacity of 90,210 L capacity (3185.73 cubic ft). Propellant mass is thus 90,210 kg. The LOX tank has a capacity of 189,270 L capacity (6674.47 cubic ft), so its propellant mass is thus 152,551.62 kg. I used a online converter for this, as well as Loru's ~10% for tank endings.

So, using the consumption rate of the infamous F-1 engine (3945 lb./s for oxidizer and 1738 lb./s for fuel, as mentioned in http://history.msfc.nasa.gov/saturn_apollo/documents/F-1_Engine.pdf, though the engines of this rocket probably won't exceed it), the burnout time (when no fuel was left in the respective tank) for the RP-1 tank was only 85 seconds and 114 seconds for the LOX tank, without throttling. It will, of course throttle, with Max Q and the fact that it has SRBs...
...

Might be going back to the drawing board (or needs more work)...

 

[ISProgram]

Revised Calculations

...bad at math

These calculations compare favorably with the
Space Shuttle ET.

LOX tank cylinder
Length: 89.73 ft. (25.35 m)
Diameter: 25.95 ft. (7.91 m)
Volume: 47,420.64 cu ft. (354,731.021 US gal; 1,342,802.98 l)
LOX propellant mass: 2,801,291.82 lb. (1,270,644.51 kg)

RP-1 tank cylinder
Length: 42.84 ft. (13.06 m)
Diameter: 25.95 ft. (7.91 m)
Volume: 22,640.14 cu ft. (169,360.0083 US gal; 641,097.37 l)
RP-1 propellant mass: 1,139,181.82 lb. (516,724.18 kg)

Please note that these calculations exclude the tank domes of both tanks. Only the purely cylindrical section has been calculated for. I will still be using Loru's ~10% suggestion for tank endings.

These calculations are based on the density of RP-1 being .806 g/cc and LOX being 1.02 g/cc, as noted by Astronautix.

Based on this, the oxidizer to fuel ratio is 2.45 (based on dividing the oxidizer mass with that of the fuel), compared to the "typical" 2.56 ratio of RP-1/LOX.
Whether or not this would be oxidizer-rich, I do not know. However, there is supposed to be excess propellant after the first stage is jettisoned, because of a secondary objective. Yes, I want to do that too. The more stages to play with in Orbiter, the better.

Lastly, how do you calculate the dry mass of a rocket stage? The current baseline for this stage is roughly 19t without propellant, similar to the SLS core stage.

As a matter of fact (SLS bootleg)...

 

[Loru]

Lastly, how do you calculate the dry mass of a rocket stage? The current baseline for this stage is roughly 19t without propellant, similar to the SLS core stage.

For dry mass I usually end up with conservative 1/10 of fuel mass for liquid stage. If you're aiming for more realism you can calculate tanks weights, support & structures and outer shell of the rocket (surface area x wall thickness x density). Solid stages are more tricky as whole stage has to withstand pressure so here I usually end up between 20 and 30% of fuel mass.

 

[Pipcard]

All I do is make my dry mass:fuel mass ratios similar to real rocket stages. Astronautix is a good resource for rocket stage data.

@Loru: The thing is, just how thick are these walls (especially if they vary from stage to stage), and what materials are they made out of?

 

[Loru]

Saturn V flight manual:

http://www.apollosaturn.com/ascom/s5flight/sec4.htm

OXIDIZER TANK

The 345,000 gallon lox tank is the structural link between the forward skirt and the intertank section. The cylindrical tank skin is stiffened by "integrally machined" T stiffeners. Ring baffles attached to the skin stiffeners stabilize the tank wall and sense to reduce lox sloshing. A cruciform baffle at the base of the tank series to reduce both slosh and vortex action. Support for four helium bottles is provided by the ring baffles. The tank is a 2219-T87 aluminum alloy cylinder with ellipsoidal upper and lower bulkheads. The skin thickness is decreased in eight steps from .254 inches at the aft section to .190 inches at the forward section.

FUEL TANK

The 216,000 gallon fuel tank (figure 4-1) provides the load carrying structural link between the thrust structure and intertank structure. The tank is cylindrical, with ellipsoidal upper and lower bulkheads. Antislosh ring baffles are located on the inside wall of the tank and antivortex cruciform baffles are located in the lower bulkhead area. Five lox ducts run from the lox tank, through the RP-I tank, and terminate at the F-l engines. The fuel tank has an exclusion riser, made of a lightweight foam material, which is bonded to the lower bulkhead of the tank to minimize unusable residual fuel. The 2219-T87 aluminum skin thickness is decreased in four steps from .193 inches at the aft section to .170 inches at the forward section.

 

[Swigert]

I was tempted to post this to many times...

 

[Phoenix]

My first thread... Planning on making a HLV addon (in the long term) to go with a fictional space program.

I've actually already designed a model with the dimensions of said HLV, though I'm having some problems calculating the fuel and oxidizer tank lengths. The aforementioned HLV uses a RP-1/LOX for its first stage and LH2/LOX for its second stages(s).

Basically, I have a question regarding the fuel to oxidizer ratio for rockets. Mainly, exactly what is that ratio? I know of the fact that rockets can have vastly different tank sizes even if they are the same diameter/length because of the different densities of preferred fuels and oxidizers (LH2/LOX or RP-1/LOX).

My knowledge on spaceflight (and programming) is very basic.

Any feedback would be appreciated...

It might be easier to think in terms of the bulk density - kilograms per cubic metre - of the overall propellant oxidizer/fuel combination, and then it's an easy calculation to find out how much your vessel can carry for a given, overall propellant tank volume. My simple rocket add-on uses this approach: [ame="http://www.orbithangar.com/searchid.php?ID=6385"]Ibis[/ame]

 

[Loru]

To quickly prototype/estimate your LV performance you can use this little tool:

http://www.silverbirdastronautics.com/LVperform.html

I found its results pretty consistent in Orbiter with ~10% error (probably due to not fully efficent autopilot)

 

[ISProgram]

All I do is make my dry mass:fuel mass ratios similar to real rocket stages. Astronautix is a good resource for rocket stage data.

The problem is, I don't think the SLS has a definitive dry mass:fuel mass ratio yet and it also uses LH2/LOX, whereupon my (SLS bootleg) uses RP-1/LOX, so it has a different ratio, to my understanding. Not to mention its design has changed a lot over the last few months, and it will probably change again.
The closest figure I have on a conclusive SLS core stage empty mass was from [ame="http://en.wikipedia.org/wiki/Space_Launch_System"]Wikipedia[/ame]...and it was 85.27 kg (188.0 lb)!!! So not trusting that (I still think Wikipedia is a valid source of information, so long as said information has citations). That previous 19t value came from a quote on a NSF Forum, and I don't remember where to find it.

OXIDIZER TANK

The 345,000 gallon lox tank is the structural link between the forward skirt and the intertank section. The cylindrical tank skin is stiffened by "integrally machined" T stiffeners. Ring baffles attached to the skin stiffeners stabilize the tank wall and sense to reduce lox sloshing. A cruciform baffle at the base of the tank series to reduce both slosh and vortex action. Support for four helium bottles is provided by the ring baffles. The tank is a 2219-T87 aluminum alloy cylinder with ellipsoidal upper and lower bulkheads. The skin thickness is decreased in eight steps from .254 inches at the aft section to .190 inches at the forward section.

...Going to have to work on that. The actual model has a 0.20 wall thickness because SketchUp has a issue with "thinness" being viewed from a distance, and it becomes almost transparent. That, though, is only for the model, which isn't terrible realistic.

What I mean:

To quickly prototype/estimate your LV performance you can use this little tool:

http://www.silverbirdastronautics.com/LVperform.html

This should prove useful, when I figure out those other variables. I don’t have a detailed flight profile yet, because that depends, currently, on the burn time of the (SLS bootleg) core stage. As well as the fact I don’t yet have a conclusive dry mass yet. I’m currently targeting a 60mT to LEO payload mass for the base configuration.

It might be easier to think in terms of the bulk density - kilograms per cubic metre - of the overall propellant oxidizer/fuel combination, and then it's an easy calculation to find out how much your vessel can carry for a given, overall propellant tank volume. My simple rocket add-on uses this approach: Ibis

Writing a vessel add-on just brought something to my mind. Is it theoretically possible to modify a existing add-on’s parameters to fit your own parameters and then fly it. A few people, I think, have done it (one guy, for example, continuously topped off the STS ET on the way into orbit, and got to the Moon with the Shuttle).


Lastly, the SRBs have been calculated:

The upper stage should be next, and most of the engine type(s) I’m making have a preliminary thrust value.

Thanks again for the support. :thankyou:

 

[ISProgram]

Testing Silverbird

OK, this should be quick...

First, a updated picture...

Second, I tested out Silverbird, with good results! The first was a control launch using the SLS, and the second was the (SLS bootleg), operating without a upper stage. There are a few errors, like the fact that payload fairing mass is still there, even without the upper stage. No doubt the payload to orbit will be lower, on the basis that the core stage will have to lift the upper stage as well with full data.

I let the pictures speak for themselves.

The SLS launch...

The (SLS bootleg) launch...

Please note that these numbers are preliminary. They will mature as I revise the data on the rocket.

 

[Loru]

. 93 MN of thrust in 1st stage?

Also I noticed you're using LH2/LOX in 1st stage.

 

[ISProgram]

. 93 MN of thrust in 1st stage?

Also I noticed you're using LH2/LOX in 1st stage.

Whoops, going to have to work on that...again. 9.3 MN it was supposed to be. Did a error in conversion, apparently.

How exactly did you come to the conclusion that I'm using LH2/LOX in the 1st stage? Is it because of the Isp or something like that in the data? I ask this because it's supposed to use RP-1/LOX and any unrealistic numbers need to be...weeded out.

Thanks, by the way, for pointing out the errors. :thumbup:

 

[ISProgram]

Tank Thickness for (SLS Bootleg)

Saturn V flight manual:

http://www.apollosaturn.com/ascom/s5flight/sec4.htm

Space Shuttle SLWT ET Analysis:

http://citeseerx.ist.psu.edu/viewdoc/download?doi=10.1.1.51.1530&rep=rep1&type=pdf

The SLWT LO2 tank shell wall, chemically milled to reduce structural weight, has a highly variable thickness distribution. In the forward ogive, the thicknesses vary from 0.080 in. to 0.157 in. in both the meridional and cir- cumferential directions. Similarly, in the aft ogive and barrel sections the thicknesses vary from 0.081 in. to 0.190 in. and from 0.140 in. to 0.385 in., respectively. In the aft dome the thicknesses vary from 0.088 in. to 0.125 in.

I can't believe how thin rocket stages are. It's almost ridiculous...I don't know how these things support their own weight (Centaur balloon tank)


Based on these figures, the aluminum alloy wall skin thickness will be .393 inches (0.01 meters) at the forward section for the core stage.

The SRBs will also be .393 (0.01 m or 10.0 mm) thick, but will be made of steel. This compares with the STS SRBs' Pre-Challenger thickness of 12.7 mm (1.0 in).

 

[Loru]

How exactly did you come to the conclusion that I'm using LH2/LOX in the 1st stage? Is it because of the Isp or something like that in the data? I ask this because it's supposed to use RP-1/LOX and any unrealistic numbers need to be...weeded out.:

ISP of 450s is close to LH2/LOX mixture. Kero/LOX is usually around 300-340

 

[ISProgram]

Okay, as it turns out, SketchUp actually has a function that can calculate volumes. This should allow me to dramatically figure out a dry mass of the stage when I get to that, as well as fairings, engines, etc. I tested it and it works.

I’m recalculating the RP-1/LOX tanks again to get the ideal (stoichiometric?) oxidizer-fuel ratio for RP-1, which is 2.56 : 1. The way I see this, about 2.56 kg or oxygen is needed to completely combust 1 kg of RP-1. My previous ratio of 2.45 is not ideal because it will leave uncombusted RP-1, unless of course, my engines is burning fuel-rich, which it is not (at the moment).

The F-1, for example, had a mixture ratio of 2.27 : 1, which means it used 2.27 kg of oxygen to combust 1 kg of RP-1, to my understanding.

…To the point, a rocket engines doesn’t have to be exactly 2.56 to be ideal, right? It could easily be something like, say, 2.36?

Because my most recent tank revisions have given me that ratio, and if that is a plausible mixture ratio for a rocket engine, then I just use that instead of going through a lot of math to calculate 2.56. Because I HATE math…

i have also been doing some calculations using Silverbird, with the decision to cross reference it with the excel worksheet from Loru to get a more "accurate" and varied performance estimate for the (SLS bootleg).

The reason? The pictures below should explain.

Unfortunately, without boosters (results not shown here), it can't launch anything at all. The 9300 kN thrust is to blame, as previous figures/estimates assumed 9300 kN. 9300 kN for 5 engines is just a little more powerful than a RS-25D, which I had based my Isp and thrust from. Of course, the RS-25 is a LH2/LOX engine, and bad one at that for this role (and SLS, I think). The reason a thrust value is not estimated yet is because the closet real engine that has a role similar to the SLS bootleg's main engines is the F-1, and it's still too big. Maybe a a RD-170...

...If you know a such a real life engine, please let me know.

Going through this thread, I founded that its following a very "unsustainable" trajectory. The pattern goes something like this:

1) I post something. 2) Someone replies. 3) I re-post, basically repeating what I just said. 4) Someone re-posts to correct/comment. 5-∞) Repeat steps 3-4 indefinitely until universe implodes.

I say this because I reread the entire post and I sounded quite stupid. I made a couple mistakes, for example, that most of anyone reading this would've noticed. Like that Isp thing.

...

This thread is going somewhere, by the way. This thread is meant to validate and estimate the realistic parameter of the (SLS bootleg), while a separate thread(?) will detail the actual making of the add-on, when I get some expertise in that area. That, being said, I don’t know how to make an add-on (meshes, scearios,etc.) yet.

Now to post this before I hesitate...