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

[ISProgram]

More Data...Again.

Once again... revised data.

Forward Skirt
Length: 13.97 ft. (4.26 m)
Diameter: 25.95 ft. (7.91 m)
Volume: 72.817 cu ft.
Mass: (Aluminum 7075-T6 density of 0.102 lb/cu in.): 12,834.43 lb. (5,821.60 kg)

LOX tank
Length: 106.72 ft. (32.53 m)
Diameter: 25.95 ft. (7.91 m)
Volume: 576.37 cu ft.
Mass: (Aluminum 2219-T87 density of 0.103 lb/cu in.): 102,584.63 lb. (46,531.60 kg)
Propellant volume: 51,939.16 cu ft. (388,531.89 US gal; 1,470,753.22 l)
LOX mass (density of 1.02 g/cc): 3,307,304.94 lb. (1,500,168.29 kg)
Cryogenic LOX mass (density of 1.141 g/cc): 3,699,642.10 lb. (1,678,129.43 kg)

Intertank
Length: 19.98 ft. (6.09 m)
Diameter: 25.95 ft. (7.91 m)
Volume: 103.544 cu ft.
Mass: (Aluminum 7075-T6 density of 0.102 lb/cu in.): 18,250.25 lb. (8,278.17 kg)

RP-1 tank
Length: 59.94 ft. (18.27 m)
Diameter: 25.95 ft. (7.91 m)
Volume: 329.58 cu ft.
Mass: (Aluminum 2219-T87 density of 0.103 lb/cu in.): 58,659.96 lb. (26,607.71 kg)
Propellant volume: 27,727.51 cu ft. (207,416.17 US gal; 785,155.64 l)
RP-1 mass (density of .806 g/cc): 1,395,163.35 lb. (632,835.45 kg)

The Entity Info function in SketchUp did most of the work for me. These calculations assume that the tank and wall thickness are all 0.02 m (0.787 in) and are uniform with regards to material (RP-1 tank, for example, just being made of 2219-T87).

 

[ISProgram]

Final Conclusion on (SLS bootleg) core stage

Final Conclusions

Total propellant mass (assuming Cryogenic LOX mass): 5,094,805.45 lb. (2,310,964.87 kg)

Preliminary dry mass (assuming only known numbers): 192,329.27 lb. (87,239.08 kg)

Preliminary wet mass (assuming only known numbers): 5,287,134.72 lb. (2,398,203.96 kg)

Estimated not-accounted-for dry mass (conservative 1/10 of propellant mass): 317,151.27 lb. (143,857.39 kg)

(NOTE: Not necessarily EXACTLY conservative 1/10 of propellant mass. About 2% of Preliminary dry mass is added for margin for estimated dry and wet mass.)

Estimated dry mass (assuming conservative 1/10 of propellant mass): 513,327.12 lb. (232,841.26 kg)

Estimated wet mass (assuming conservative 1/10 of propellant mass): 5,608,132.57 lb. (2,543,806.14 kg)

I kn ow it's a lot of numbers, but only the last two are relevant. That is the final number on....well the data says it itself, so I don't need to say anything. This is still a number subject to change, though only if anyone points any thing out. Those last two numbers do/do not account for the five engines, based of course on that 1/10 ratio suggested by Loru.

@Loru: Does that ratio actually account for (any) engines?

Next, since I have the mass of the stage, I can calculate the thrust it will need for orbit. Expected thrust is to be enough to get a single core flying on its own, or at least in a high enough altitude for staging or Ares-IX style mission, due in part because of a mission I did.

 

[Loru]

... Those last two numbers do/do not account for the five engines, based of course on that 1/10 ratio suggested by Loru.

@Loru: Does that ratio actually account for (any) engines?

1/10 of fuel mass for liquid stage dry mass is an estimate incuding entire stage (structure + engine + tanks) Also it's just estimate - some rockets do more, some do less. Look at various stages (both historical and present) that are using similar fuels to get more results and you'll see how figures vary from stage to stage. Atlas V and Saturn V are good examples here (both kero/lox).

 

[ISProgram]

Sorry guys, this again...

Again, the core stage is getting revised. I found a whole new set of structural and mass problems, among other things. I will not get into that yet, other than to mention that I am possibly considering swapping the tanks, so that the RP-1 tank is on top of the LOX tank. Not sure how this will affect the rocket, other than the fact that it may complicate launch infrastructure.

What I mean:

This was done after I realized that the SRB attach points were right next to the LOX tank. Looking at a Space Shuttle ET diagram (or any SDLV), it should be noted that a support structure is between the two attach points, through the intertank.

Since I don't think any other RP-1/LOX rocket has a configuration like this, it must not be a good one. Anyone have any opinions on this?

EDIT: I have decided that the tank configuration will remain as it is. Reason being that the Ariane 5 does not possess such a support structure, despite having large SRBs.

 

[ISProgram]

After a good look at some realistic RP-1/LOX rocket stages, I have subsequently decided to lower the dry mass estimate for the core stage. It is now a preliminary 330,000 lb., down from 513,327.12 lb. and based in part because I noticed that the (SLS bootleg) core stage and the Saturn V S-IC stage have many similar specifications. It will probably change again, since my mind obviously is not made up

Saturn V data based from this source:
(Saturn V vs. previous SLS bootleg data)
Loaded Mass: 4,881,000 lb. vs. 5,608,132.57 lb.
Dry Mass: 303,000 lb. vs. 192,329.27 lb.
LOX mass: 3,178,000 lb. vs. 3,699,642.10 lb.
RP-1 mass: 1,400,000 lb. vs. 1,395,163.35 lb.

Those numbers are pretty close to each other relatively speaking, even when accounting for the Saturn V’s 1:2.27 mixture ratio (vs. the SLS bootleg’s current 1:2.65) or the fact that the F-1 engines account for a large fraction of the mass of the S-IC, whereupon the SLS bootleg does not.

The S-IC has a dry mass of 210,500 lb. without the 5x F-1 engines, which gives it a fuel mass/dry mass ratio of 22.70 (16.10 with the 5x F-1s). I will be targeting this ratio for my SLS core stage, since they are similar in the aforementioned data. That is because I don’t have definitive engine mass data yet, and the fact that calculating the core stage without any engines and then calculating engine mass, etc., would be easier for me, since the SLS bootleg estimates are entirely based off the propellant mass data and how that correlates to the rest of the stage. Since the current propellant mass data turned out to be erroneous, I have a real problem at the moment. So…

Previous numbers had assumed the tank’s full capacity and that it was devoid of any internal structures (it was a shell), when the fact is that the tanks are not completely filled (didn’t know about this) and that that there are internal stuff like the slosh baffles, etc.

@Anyone: Anyone know how to what percent of a “shell” tank is used by internal structures? Also, should I honestly be trying THIS hard to calculate all this information, or would an “honest” estimate suffice? This is becoming tedious, that’s why I’m asking. Not to mention I want to get to Part 2 (meshes and scenarios) soon.

Started that already, by the way. Don't quite have the hang of Anim8or yet:

 

[ISProgram]

Latest (and hopefully final) data on the core stage of the SLS bootleg.

Overall
Height: 196.03 ft. (59.75 m)
Diameter: 25.95 ft. (7.91 m)
Total propellant mass: 5,073,500 lb. (2,301,300.88 kg)
Dry mass: 292,822.20 lb. (132,821.91 kg)
Wet mass: 5,366,322.20 lb. (2,434,122.80 kg)
Engine(s): 5x A-3 (F-X)
Thrust: 6,620,000 pounds-force (29,447 kN)
Burn time: 480 s (?)

Some values will need editing...

 

[ISProgram]

Change of Plans

Well, I decided that until I can resolve the problem with the Hermes C (SLS bootleg), it has indefinitely been put on hold. The core problem with the rocket was the unfortunate fact that I don’t have any accurate data for a rocket of its type, since there is no RP-1/LOX rocket of its size. Only the S-IC comes close and it has its own sets of problems.

So, in the meantime, I will be making an addon for another IPSA rocket, the Aquarius 1-IP. I already calculated all the data, so there will be none of this going-back-and-forth like I was doing the past few months.

At the moment, it is capable of 43,319 kg into a 185 km x 185 km orbit with an inclination of 45 deg, launched from Cape Canaveral. This information comes from Silverbird.

If the payload seems a tad bit high (more than the Delta IV Heavy), I thought that too, and even considered the possibility of an error, until I realized Aquarius was essentially an Atlas Phase II. So that solved that.

The work on that aforementioned addon will begin shortly once I figure out how to make an addon in its entirety. Saying this, I’m asking if anyone can help give me the basics or a link to a helpful resource.




UPDATE: Some payload mass numbers (not including attach fittings)

(NOTE: Payload mass to given orbits is from Cape Canaveral. Inclination and/or Declination is 28.5 deg)

Mass to LEO (circular orbit of 160 - 2000 km): 46,760 – 9,661 kg

Mass to MEO (orbit w/ apogee of 20,200 km and perigee of 185 km): 20,185 kg

Mass to GTO (orbit w/ apogee of 35,786 - 42,164 km and perigee of 185 km): 18,116 – 17,413 kg

Mass to Venus (Hyperbolic C3 of 11 and orbit w/ perigee of 185 km): 9,967 kg

Mass to Mars (Hyperbolic C3 of 17 and orbit w/ perigee of 185 km): 8,632 kg


Also getting something actually done:

Admittedly, I "hijacked" the a .msh file from the [ame="http://orbithangar.com/searchid.php?ID=6438"]M-II + Negi-5 launch vehicles v0.2.1[/ame], so I apologize to Pipcard in advance.