Calculating the dry mass of a 5m+ diameter stage.

[ISProgram]

50mt to LEO!

Okay, so I've done it, the rocket now gets 50,000 kg to LEO!

The first stage still had some margin on its propellant mass, 0.03%, which is now gone. The LR-60 dry mass has been reduced to 500 kg. Since the LR-60 is smaller than either the L-1 and its future replacement engine, the interstage could lose weight. Payload fairing mass and jettison time is unchanged.

Here are the numbers:

First Stage
Empty Mass/Inert: 37,863 kg (83,474 lb)
Thrust: 12,040 kN (2,706,700 lbf), 70% of full thrust
Propellant Mass: 809,326 kg (1,784,258 lb)

Second Stage
Empty Mass/Inert: 8,894 kg (19,608 lb)
Thrust: 1,160 kN (260,700 lbf)
Propellant Mass: 79,296 kg (174,817 lb)

Interstage
Mass: 1,096 kg (2,416 lb)

Silverbird says the rocket can get 50,127 kg to a 160 km circular orbit with a inclination of 28.7, launching from Cape Canaveral. Admittedly, I would've preferred it to be 185 x 185, but I've exhausted practically all my margin, and rockets are tricky things.

The first stage has a delta-V of 5652.17 m/s and the second stage has one of 3931.08 m/s. This does not include payload fairing mass on the second stage, because the fairing will either be jettison during first stage burn or shortly after second stage ignition.

Complete delta-V is thus 9583.25 m/s, so this redundantly confirms that it has a 50,000 kg payload capacity.

Success!

 

[ISProgram]

50+mt to 185 LEO

Okay, working on my Orbiter repertoire right now, so I'm learning all I can about Velcro and Spacecraft3, all those meta-addons, since I will try to make a addon of this relatively soon.

The very last margin on the rocket was unused volume because the tank skin volume for the developmental model was still a little big by about a 1mm. The extra fuel the updated model gave wasn't a lot, BUT was enough to get 50,000+ kg to a circular 185 LEO.

First Stage
Empty Mass/Inert: 37,863 kg (83,474 lb)
Thrust: 12,040 kN (2,706,700 lbf), 70% of full thrust
Propellant Mass: 809,942 kg (1,785,616 lb)

Second Stage
Empty Mass/Inert: 8,894 kg (19,608 lb)
Thrust: 1,160 kN (260,700 lbf)
Propellant Mass: 89,840 kg (198,063 lb)

Silverbird says the rocket can get 51,071 kg to a 185 km circular orbit with a inclination of 28.7, launching from Cape Canaveral.

Delta-V is 9551.42, with 5,613.53 m/s and 3,937.89 m/s coming from the first and second stage, respectively.

This should/might be the penultimate post dedicated to the first rocket's design. The only uncertainty now is the interstage length, which will depend on the length of the second stage, which already has a Not To Exceed value.

My personal opinion towards designing rockets has changed radically as a result of this thread, so thanks to all involved.

BTW, the rocket's called the Aquarius B1 (Block 1).

EDIT: I made a unknown error in my estimates. Propellant mass for second stage is actually 80,946 kg (178,455 lb), while the delta-V for first stage is 5629.29 m/s, and the second stage is 3,985.79 m/s. Silverbird estimate to LEO is now just 50,283 kg.

 

[Phil Smith]

Nice! waiting for the addon))

My personal opinion towards designing rockets has changed radically as a result of this thread, so thanks to all involved.

Great to hear! You're welcome!

BTW, the rocket's called the Aquarius B1 (Block 1).

It reminds me one song :lol: :

 

[RGClark]

Thanks again for the help! 2.19 nozzle diameter?! :jawdrops: How am I going to get FOUR of those on a 5.75m rocket?! It's Frankenstein's monster at this point!

And I should supply more detail in the future, that 17,290 kN is its vacuum thrust, since these numbers were made to be fed into Silverbird. Sorry about that.

The theory of circle packing is well developed for the case of packing within a larger circle:

http://en.m.wikipedia.org/wiki/Circle_packing_in_a_circle

For 4 equal circles, the diameter needed is 1 + sqrt(2) times that of the 4 small circles, about 2.414 times. So for 2.19 diameter the larger diameter would need to be 2.414*2.19 = 5.287.

Bob Clark

 

[ISProgram]

The theory of circle packing is well developed for the case of packing within a larger circle:

http://en.m.wikipedia.org/wiki/Circle_packing_in_a_circle

For 4 equal circles, the diameter needed is 1 + sqrt(2) times that of the 4 small circles, about 2.414 times. So for 2.19 diameter the larger diameter would need to be 2.414*2.19 = 5.287.

Bob Clark

Yep, the nozzle diameter for the four-chambered first stage engine is 2.11m. The core itself is 5.75 m.

Now, I've decided to use Velcro for the add-on, it seems easier to program a vessel with than spacecraft3. The only problem I have is that I don't particularly like the exhaust textures.

I've run into a slight problem, not directly related to the add-on progress. I've decided to add another launcher with the Aquarius.

I decided to try doing this one's calculation because I thought I knew enough to do it, but an issue popped up. Namely, the tank skin thickness value.

The rocket's first and second stage are both 3.91m in diameter. The first stage is identical in height to the Atlas V CCB. I used a tank skin value of 5mm, and the first stage engine used is 5,419 kg. Yet it's only 13,226 kg overall. The Atlas V CCB is 3.81m (same height) and weighs 21,054 kg. The Zenit-II first stage is only slightly taller (same diameter), and weighs 27,564 kg.

It seems to me that my rocket's a bit....light. Any suggestions? Or should I just go with a 6mm tank skin thickness?

 

[kocmolyf]

It seems to me that my rocket's a bit....light. Any suggestions? Or should I just go with a 6mm tank skin thickness?

There's a lot more to a rocket than a tank and some engines (and even those are complicated). When real rocket scientists are doing trade studies, they will typically employ what are called "mass-estimating relations" (empirical formulas with input parameters like height and volume) to do it. These relations help account for all the real engineering that will have to be done beyond the basic e.g. tank wall thickness.

Short version: it's more realistic to estimate based on existing examples than to calculate from first principles, at least for us amateurs. Your rocket probably is too light.

 

[ISProgram]

There's a lot more to a rocket than a tank and some engines (and even those are complicated). When real rocket scientists are doing trade studies, they will typically employ what are called "mass-estimating relations" (empirical formulas with input parameters like height and volume) to do it. These relations help account for all the real engineering that will have to be done beyond the basic e.g. tank wall thickness.

Short version: it's more realistic to estimate based on existing examples than to calculate from first principles, at least for us amateurs. Your rocket probably is too light.

I was expecting around 20,000-25,000 kg for the rocket. That's why I'm surprised that it's that low.

I'm aware that rockets are not just tanks and engines. But I was using a equation/formula that Phil Smith provided. There's that formula, with the values as they applied to my rocket:

Cylindrical tank mass = Stage diameter * Pi * Skin thickness * Stage height * Aluminum density

RP-1 Cylindrical tank mass = 3.91 * 3.14 * 0.005 * 7.08 * 2840 = 1234.3206864 kg

LOX Cylindrical tank mass = 3.91 * 3.14 * 0.005 * 14.73 * 2840 = 2568.0146484 kg

Total mass of RP-1/LOX tanks = (Bulkhead*4) + RP-1 cylindrical tank mass + LOX cylindrical tank mass

Total mass of RP-1/LOX tanks = (268.096*4) = 1072.384 + 1234.3206864 + 2568.0146484 = 4874.7193348 kg.

Skin mass = Stage diameter * Pi * Skin thickness * Skin height * Aluminum density

Forward skirt mass = 3.91 * 3.14 * 0.005 * 0.86 * 2810 = 148.3478242 kg

Interstage mass = 3.91 * 3.14 * 0.005 * 3.67 * 2810 = 633.0657149 kg

Aft skirt mass = 3.91 * 3.14 * 0.005 * 4.35 * 2810 = 750.3639945 kg
750.3639945 + Bulkhead area * Skin thickness * Aluminum density
750.3639945 + 11.87 * 0.005 * 2810 = 917.1374945 kg

Skirt mass = 148.3478242 + 633.0657149 + 917.1374945 = 1698.5510336 kg

Other mass = (RP-1/LOX tank + Forward/aft skirt) * 0.15

Other mass = (4874.7193348 + 1698.5510336) * 0.15 = 985.99055526 kg.

Total stage dry mass = RP-1/LOX tank mass + Skirt mass + Other + Engine mass + Thrust structure

Total stage dry mass = 4874.7193348 + 1698.5510336 + 985.99055526 + 5419 + 247 = 13225.26092366 kg.

I wold assume the "Other" variable would constitute anything not a tank or engine.

Also, I realized that I might use different alloys, so that might explain one cause for the discrepancy.

 

[kocmolyf]

I'm aware that rockets are not just tanks and engines.

Sorry, I only meant to suggest that there's more to a rocket than its skin. For example, here's the inside of a Saturn V S-1C tank (I believe this is the fuel tank):

This kind of thing is why the simple model of a cylindrical shell doesn't hold up very well.

 

[ISProgram]

Sorry, I only meant to suggest that there's more to a rocket than its skin. For example, here's the inside of a Saturn V S-1C tank (I believe this is the fuel tank):

This kind of thing is why the simple model of a cylindrical shell doesn't hold up very well.

For the Aquarius rocket, which I was doing earlier, I asked the same question about internal structures. The tank skin measurement, as it pertained to the calculation I did was 7-5mm, though Phil Smith said that an "actual", real life measurement of the tank skin would have been 4-5mm and that the extra was there to account for internal structures. I did that for this rocket as well.

He specifically mentioned stringers (I would've liked isogrid, but don't know how to calculate that, since it's so diverse in terms of application).

Incidentally, increasing the skin thickness by 1mm only increased by a little over 1,000 kg; it's 14,737 kg right now.

---------- Post added at 10:55 PM ---------- Previous post was at 10:53 PM ----------

Can anyone give me tips on how they calculate rocket dry mass? :facepalm:

Until then, I just continue with the addon prep.

 

[ISProgram]

I'm going to have to apologize before I say this, that the help the O-F community has given me may not be put to good use.

First, I'm still making the addon; that hasn't changed. Not the bad news.

The bad news is that preliminary recalculations of the Aquarius B1 have the first stage weighing around 42,641 kg and with a reduced fuel load. These changes will also affect the second stage. In all, I don't expect to get past 40mT to LEO now. Triple core configuration will probably be 80mT, propellant crossfeed about 110mT.

It seems the earlier estimate by Phil Smith is the more realistic accurate value, then (44,941 kg).

---------- Post added 11-07-14 at 03:19 AM ---------- Previous post was 11-06-14 at 10:13 PM ----------

Okay, I just remembered that I learned how to make a speadsheet in class today.

So I made a (overly complicated) spreadsheet to calculate the rocket's dry mass with the equation. Came in to about 43,225 kg 46,196 kg.

Second stage pending. The spreadsheet's in Google Drive, if anyone knows what that is. I didn't two weeks ago. :lol:

---------- Post added at 03:58 AM ---------- Previous post was at 03:19 AM ----------

The 3.91m launch vehicle has a dry mass of 14,475 kg. Expecting around 23,483 kg. :blink:

---------- Post added at 08:03 PM ---------- Previous post was at 03:58 AM ----------

The rocket has a established payload of 45,826 kg with earlier orbital parameters. The addon will continue, minor edits to .cfg files.

Here's the spreadsheet:

 

[ISProgram]

Okay, have been trying to get payload back up to 50mT following all the stuff that's happened. So far, only .cfg for the rocket and some preliminary .scn file have actually been worked on. Probably going to have to change the .cfg afterwards again.

So, I decided to go try to put tank barrels into the rocket (didn't do this before) since the upper and second stage share tooling techniques. This resulted in a lengthening of the S2's LH2 tank to that of the S1's RP-1 tank. The LOX compartment was remodeled so that only one bulkhead type is in use for the entire vehicle, with the result being that the upper stage has lengthened by <2 meters and it has a estimated >2,000 kg more propellant.

The first stage also has an increased prop load, though I didn't lengthen its tanks as much. The bulkheads themselves have been changed from a 70% ellipsoid to a 75% one. Intertank and aft skirt have been shortened.

Fairly confident that will get >50mT to LEO.

@ Phil Smith: I've been trying to spread-sheetify the calculation used for bulkhead thickness in MS Excel for a while now, but I can't get past arcsin (which I'm starting to doubt has a numerical designation/equation that can be implemented at all). Any advice?

---------- Post added 11-18-14 at 12:45 AM ---------- Previous post was 11-17-14 at 11:22 PM ----------

Current estimates have 49,954 kg to LEO. :facepalm: A tad bit more...

Increasing the thrust level to 12,400 kN (so that the LV has a liftoff T:W of 1.3 as suggested earlier by kocmolyf) gives a 50,252 kg LEO payload.

First stage has delta-v of 5,316.74 m/s, second stage 4,356.80 m/s. Total delta-v is 9673.54 m/s.

So back on track, AqrB1 is still a 50mT LEO rocket.

 

[Phil Smith]

Hey ISProgram!
Sorry mate, it's been tooooo long, almost a year)) I got a lot of work...
How is your Aquarius B1 vehicle? is it still in developing stage?
If you got some questions about rocket structure / engine design - let me know.
Better hit me with email - philsmith29990 "at" yahoo.com , cause I'm not a daily visitor of the orbiter forum. Unfortunately.

..here's the inside of a Saturn V S-1C tank (I believe this is the fuel tank):

...

yep, but it's a LOX tank - anyway - this photo is a great example of how complicated a humble rocket tank can be (because of its size)... This big rings are slosh baffles, the huge structure in front of the camera (at the bottom of the tank) is a vortex baffle assembly. Two technicians are hanging around 4 Cold helium cylindrical tanks (they are parts of fuel tank pressurization system), and if you look closer you'll see wall T-shaped stringers are penetrating whole tank structure to withstand all flight & manufacturing loads, including axial / shear forces and bending moments (about 2-2.5" height with 6-8" spacing).

Actually I can discuss this pic and all this stuff all day long, so let's stop right now))

Cheers!

PS. Sorry for bumping the old thread..

 

[N_Molson]

Usually, the "10% thumb rule" isn't too far from the truth, at least for conventional liquid-fueled multi-staged rockets.

If you have 800,000 kg of propellant, a dry mass of 80,000 kg isn't crazy.

 

[ISProgram]

Hey ISProgram!
Sorry mate, it's been tooooo long, almost a year)) I got a lot of work...
How is your Aquarius B1 vehicle? is it still in developing stage?
If you got some questions about rocket structure / engine design - let me know.
Better hit me with email - philsmith29990 "at" yahoo.com , cause I'm not a daily visitor of the orbiter forum. Unfortunately.



yep, but it's a LOX tank - anyway - this photo is a great example of how complicated a humble rocket tank can be (because of its size)... This big rings are slosh baffles, the huge structure in front of the camera (at the bottom of the tank) is a vortex baffle assembly. Two technicians are hanging around 4 Cold helium cylindrical tanks (they are parts of fuel tank pressurization system), and if you look closer you'll see wall T-shaped stringers are penetrating whole tank structure to withstand all flight & manufacturing loads, including axial / shear forces and bending moments (about 2-2.5" height with 6-8" spacing).

Actually I can discuss this pic and all this stuff all day long, so let's stop right now))

Cheers!

PS. Sorry for bumping the old thread..

Nice to hear from you again! It has been too long...

As far as developing the actual rocket, the design is fully finalized; the last major edit was roughly six months ago (see the bottom of this post for Silverbird values). If anyone feels like picking it apart, feel free. :thumbup:

I have started work on the add-on, there's a thread for that now, but work has slowed on it too for a bit, for various reasons, the most prominent being that I was using a "learn as you work" approach, which was rather detrimental. Work is still going on, albeit at a much slower pace.

Aquarius B1-C Silverbird Specifications:

(I use the "C" to denote this as the third major revision it's gone through since conception)

1st Stage
Empty Mass/Inert: 45,135 kg
Thrust: 12,400 kN
Propellant Mass: 821,840 kg
Isp: 338 s in vacuum

2nd Stage
Empty Mass/Inert: 10,930 kg
Thrust: 1,160 kN
Propellant Mass: 96,228 kg
Isp: 460 470 s in vacuum

Payload Fairing
Mass: 2,600 kg
Jettison: 267 s

Any questions, I'll try to answer them.



---------- Post added 11-01-15 at 11:56 AM ---------- Previous post was 10-31-15 at 07:53 PM ----------

I feel that while the actual Aquarius B1-C launch system is more or less at a CDR stage of development, the ground support logistics and infrastructure are still very much at PDR level. 2nd stage Isp is actually 470 seconds, apologies.

 

[kerlix]

I apologise for bumping an old thread, but has anyone creates a spreadsheet or program that has all of these equations and problems in it so that one could just insert the variables and get the information?

My math isn't the greatest and I have little to no programming skills. I wouldn't even know where to begin to turn this into a spreadsheet.

I know all the information is here. How difficult do you think it would be for someone like me to turn this into a program?

And what coding language would you suggest? I was thinking C# but I may very well be wrong.

Sent from my PH-1 using Tapatalk

 

[escorpio]

Hi, maybe this spreadsheet will be useful, assume that the dry mass is 10% of the mass of fuel and oxidant

https://www.orbithangar.com/showAddon.php?id=b9e6b928-51c9-470a-b10e-405c05694e30