kbd512 Postings

tahanson43206 · 2025-04-09 17:11:58

For kbd512...

Thanks for looking at the specs for the optical plane.

Your comment revealed an (understandable) impression that the AI picked 1000 as the solar flux I wanted to design for. That is not correct. I told it from the beginning that I wanted to use 1000. If we design for 1000, and get more, then the system will perform better. On the other hand, if we plan for 1000 and fly out to a distance where 1000 is normal, then the system will perform as designed.

***
Thanks for additional details about how metals might help at some point in the project. I've decided to just go with the recommendation to use carbon, based in part on the great number of examples you provided. In my next session with ChatGPT4o, we'll dive into the challenge of heating the hydrogen through a 240 meter long carbon tube. At this point I don't know what diameter to start with or what increases to make along the 240 meters, but I'm hoping we can figure that out as we go along.

The optical input from the wings will be 6000+ watts per 38 mm, and that input will be supplied for the entire length of the pipe.

ChatGPT4o seems to think that 40 MW can heat 1 kg/s of hydrogen from 300 K to 3000 K, so tomorrow I hope to see how that would happen in the long heating tube.

I'll also be interested to see what pressure ChatGPT4o comes up with as we proceed. GW was looking for something in the range of 3.5 to 7 bar, and apparently if we can provide that he can deliver .93 force-tons (point nine three).

(th)

tahanson43206 · 2025-04-12 16:35:28

For kbd512 re graphite as a candidate material for Heat Pipe...

GW wrote about graphite in the context of heat shields and rocket nozzles.

My guess is he is not up to date with all the new materials you have found that feature carbon.

Graphite is just one form of carbon.

If you have time, please prepare a review of carbon based materials for Sunday evening.

https://newmars.com/forums/viewtopic.ph … 41#p231041

Graphite is both a poor insulator and a poor conductor of heat.

(th)

tahanson43206 · 2025-04-14 19:48:54

For kbd512 re Optical Plane vessel...

There is a need to be able to focus either the optical devices individually, or the entire wing to stay in the Sun's view as much as possible.

I am starting to work on navigation, and we may not be able to guarantee 100% perpendicular orientation for the space tug when attached to a payload.

Can you help by doing a bit of research to find out how much flexing fiber can stand? We're not talking about much but most fiber never moves after it is installed.

(th)

tahanson43206 · 2025-04-15 09:32:07

For kbd512 re Optical Plane vessel...

I'm working on a program to combine heat pipe and nozzle.

I looked for carbon pipe but only found carbon steel.

If you can spare a moment, please see if you can find pipe of about 25 cm diameter or 9 inches or a bit more.

The heat pipe needs to be larger than the nozzle throat, but this seems to be a chicken-egg problem so I thought a reasonable approach would be to see what size pipe is available.

(th)

kbd512 · 2025-04-16 04:32:05

tahanson43206,

Graphite heat exchanger tubing and shells are available in a very wide variety of sizes and shapes, because it's a sintered and then hot-pressed material formed using molds or 3D printing. I don't know think there are any Graphite tubes 240m long, but there are certainly 5m to 10m long tubes with flanges for connection to additional lengths of Graphite tubing. Heck, we even make Graphite and RCC bolts / nuts / studs / gaskets for connecting flanged piping.

If you require a 240m long heat pipe, that would imply 24 of the 10m long tubes, and possibly some sort of exterior support structure using stranded Carbon Fiber acting sort of like the suspension bridge cable-stays, which can natively survive high temperatures since they're made from pure Carbon as well.

complex pure Graphite components such as this jet mixer are also possible:

The components in some of these chemical processing plants are of considerable size because they run the Sulfuric acid through one set of pipes while they run water outside the pipe to absorb heat from the chemical reaction taking place inside the smaller diameter Graphite tubing. These chemical processes run at modest pressures but can and sometimes do achieve high temperatures.

tahanson43206 · 2025-04-16 06:11:19

For kbd512 re #255

Thanks for additional information about graphite tubes and related products!

Thanks too, for thinking ahead to mounting the heat pipe inside the cylinder of optical output devices.. The forces at work during flight will be primarily experienced along the length of the pipe, but keeping the pipe centered in the heating cavity is important, and your suggestion of a tension network instead of a compression system should save weight.

The hint I pick up from your post is to ask Google to look for chemical plant suppliers and for graphite tubing.

As reported earlier, the diameter of the heat pipe appears to be a significant factor in design of the nozzle, so I'm hoping to find a commercially available pipe suitable for this application.

(th)

tahanson43206 · 2025-04-19 19:22:03

for kbd512 re Registration message....

I put a reminder in Housekeeping, as you suggested.

I ** think ** that since you are an Administrator, you may have the ability to update that message in the database.

SpaceNut would ordinarily take care of something like that, but he's been out of touch since his wife died, and he started the house remodel.

(th)

tahanson43206 · 2025-04-20 19:58:30

For kbd512 ....

First, thanks for updating the Registration Information Line !!!

Second, re https://newmars.com/forums/viewtopic.ph … 40#p231240

Holy Moley!

Third... I was able to take a picture of the Merlin engine outline:
https://newmars.com/forums/viewtopic.ph … 39#p231239

(th)

kbd512 · 2025-04-20 21:55:16

Blackbody Definition
A blackbody is a theoretical object that absorbs 100% of the radiation that hits it. Therefore it reflects no radiation and appears perfectly black. In practice no material has been found to absorb all incoming radiation, but carbon in its graphite form absorbs all but about 3%.

Stefan-Boltzmann Law
P = σAT4
P = Power (energy per unit of time)
σ = Stefan-Boltzmann constant (5.67 x 10^-8 W/m^2K^4)
A = surface area of the blackbody (in square meters)
T = temperature (in Kelvin)

A = 1m^2
T = 3,000K
P = (5.67 x 10^-8 W/m^2·K^4) * (1m^2) * (3,000K)^4
P = 0.0000000567 * 1 * 81,000,000,000,000
P = 4,592,700 Watts per square meter, or 459.27W/cm^2

40,000,000W = total collected input power available from the solar concentrating array for heating of the Hydrogen propellant

40,000,000W / 4,592,700W/m^2 = 8.709m^2 <- Maximum receiver surface area which can be heated to 3,000K by the solar concentrator array

For this Graphite heat pipe (more of a soda straw) to be 240m in length, it needs to be 11.54mm in diameter. On the bright side, it could be flexible enough to bend like a fishing rod for easy storage / transport to LEO. On the downside, coupling the power into something that small could present a bit of a challenge. However, I think I finally understand what you're trying to tell me about why it has to be built that way. It's related to getting the power out of the collector array and focused onto the receiver tube assembly, as I understand it. I'm not trying to argue, merely understand what you're telling me. If that's how it has to be built to keep the weight of the collector array in check, as well as the complexity and thus cost of the tech used, then so be it. I can envision some structural advantages to the 240m tube vs a much more compact receiver. After all, that gigantic solar concentrator array has to be connected to something.

The emissivity of Graphite is not equal to 1 (Graphite is not a perfect black body), so we would need to modify the equation used to arrive at the actual Watts of power input required to heat the receiver tube assembly to 3,000K. However, this places an upper limit on the surface area and volume.

What are we connecting the LH2 tanks to, so as to keep them well away from the receiver tube?

We seem to have a rather complex vehicle geometry to keep the imparted forces low and the structures light, but I'm trusting your judgement on this arrangement so as to ensure we remain within tolerable mass limits because you have more sophisticated tools at your disposal than I do, specifically the AI and OpenFOAM simulation tools.

tahanson43206 · 2025-04-21 10:50:44

For kbd512 re #259

Thanks for this post, with what appears (at first scan) to be a helpful summary of data about the heat tube.

It will be a bit of a while before I can get back to it, but I'll put a search term here so it's easy to find: SearchTerm:Graphite

Void just posted about 3D printing of carbon structures. I don't know if it applies to the Optical Plane vessel, but the idea of 3D printing the entire heat tube in one operation is attractive.

https://newmars.com/forums/viewtopic.ph … 46#p231246

Update later....

If it is possible to 3D print a tube 240 meters long on Earth, then it should be possible in LEO. That would be a ** very ** nice way to solve the assembly-on-orbit problem. If it is possible to 3D print a tube, then it should be possible to print the spars and ribs of the Optical Plane wings, if they can be made of carbon, which seems like a reasonable idea to me.

***
In post #259 (if I am remembering correctly from my first scan) you brought up the important point of the keel design, to keep the propellant away from the hot tube at the same time as keeping the propellant away from the Sun. One advantage of the ship design is that there will be plenty of electricity available to keep the propellant cool while preparing for thrust events.

I suspect that some metal will be required for the keel, because the engine thrust is going to be carried up to the payload from the nozzle, according to GW's latest iteration of the design. If you have not been following along, the nozzle cannot swivel, so the most attractive option right now is to swivel the docking adapter.

(th)

tahanson43206 · 2025-04-28 06:13:54

For kbd512 re ADEPT and NIAD

https://newmars.com/forums/viewtopic.ph … 29#p231329

Thanks for this addition to the heat shield topic!

I added an Index entry in Post #2 of the topic.

And thanks again for thinking about design of the heat pipe for the Optical Plane vessel.

Since nothing like this has ever been attempted before, surprises await.

While carbon is the leading candidate (with a liner inside) it may not win the competition.

(th)

tahanson43206 · 2025-04-29 09:14:47

For kbd512 re posting in Heat Shields

https://newmars.com/forums/viewtopic.ph … 54#p231354

Thanks for this addition to the topic.

Your point about thickness of the coating reminds me we (or for sure ** I ** ) don't know the thickness of the heat shield used on Apollo. I am interested it seeing if the successful Apollo heat shield design can be brought forward using modern technology, so humans do not have to squeeze goop into tiny cells of fiberglass using grease guns.

If you can (somehow) discover the thickness of the heat shields that were glued to the metal surface of the Apollo capsules, we would have available guidance for design of identical heat shields made with modern techniques.

It would also be valuable to know the width of the hexagonal cells in the successful Apollo heat shields.

I assume that the height of the honeycomb was identical to the thickness of the ablative material, but I don't actually ** know ** that.

There is a business opportunity for a company that can make heat shields for the deep space shipping market, until improved propulsion techniques make use of heat shields unnecessary.

The only proven heat shield method for deep space returns with humans as cargo is the Apollo method.

I wonder what the Chinese are using for their heat shields.

(th)

tahanson43206 · 2025-04-30 19:38:34

For kbd512 re Optical Plane vessel...

ChatGPT4o has determined that the mesh we are working with has a diameter of .4 meters (40 centimeters) at the intake.

We are starting work on the heat pipe tomorrow. 20 centimeters is the diameter that came out of another work session, in which we were trying to anticipate the behavior of hydrogen in heating up over 240 meters. I'm hoping you can give a few minutes thought to the problem of feeding 6325 outputs onto the heat pipe every 38 millimeters. We can resize the mesh to match whatever diameter pipe is best. At this point, we have not much to go on, except for that 20 centimeters. That diameter matched up with 5 bar, 3000 Kelvin and flow rate of 1 kg/s.

For now, unless advised otherwise, I'll stick with 20 cm.

(th)

tahanson43206 · 2025-05-07 11:41:38

For kbd512 re post in plastic cars topic:

https://newmars.com/forums/viewtopic.ph … 93#p231493

Outstanding find! Thanks!

Question: With such a strong start, can this venture source all components locally? That would really be something!

(th)

tahanson43206 · 2025-05-15 06:08:58

For kbd512 re Mars use of hydrazine ....https://newmars.com/forums/viewtopic.php?pid=231646#p231646

It seems to me (if I remember correctly) "The Martian" used storable propellant as a plot device.

The stranded astronaut burned fuel to make water. And, (come to think of it) the vehicle that he rode to orbit must have had the same propellant.

If my memory is correct, that would appear to support your argument.

Update:

In "The Martian," both the book and movie utilize hydrazine as the main rocket fuel. Hydrazine is a monopropellant, meaning it can burn by itself without needing a separate oxidizer. In the movie, Mark Watney uses a catalyst (likely iridium) to decompose hydrazine, extracting hydrogen to produce water, according to LabXchange.

I missed that detail about how the hydrazine was decomposed... I thought it was by burning, so am glad to have this correction.

Follow up ... OK ... Google appears to have gotten the detail about the catalyst but missed the burn to make water... A web site interviewed Dr. Zubrin ...

https://www.labxchange.org/library/item … 6e5:html:1

Synthesizing Water
Mark makes use of hydrazine from the rocket fuel and goes on to dissociate it into two constituent elements, i.e. nitrogen and hydrogen. Subsequently, he used the hydrogen and burned it with cabin oxygen to make water. This way of making water will definitely work, so the movie got that detail right too!

Follow Up #2... Are there any risks to a human crew, in flying with tanks of hydrazine sufficient to return from Mars?

I'm wondering if those risks were greater than NASA risk tolerance?

(th)

kbd512 · 2025-05-15 09:35:02

tahanson43206,

Every crewed American spacecraft since Gemini, as well as nearly every robotic probe, has used some kind of Hydrazine-based propellant for RCS, mostly NTO/MMH, but sometimes Aerozine-50 or UDMH. I think the Mercury program used something weird like Hydrogen Peroxide monopropellant decomposition. Dragon, Orion, Starliner, and Dream Chaser all use NTO/MMH. Both stages of the Lunar Excursion Module used Hydrazine-based propellants as well. All other crewed spacecraft used by all other space-faring nations use Hydrazine-based monopropellants and bipropellants as well.

There's no serious effort to replace these oxidizers and fuels, although certain Hydrazine substitutes such as Hydroxyl-Ammonium Nitrate (HAN) or similar chemicals can serve as monopropellants with improved Specific Impulse over Hydrazines and greatly improved Density Impulse. For equal Total Impulse, HAN only requires half as much propellant tank volume as MMH. Accidentally freezing HAN does no damage to the propellant or tankage, unlike Hydrazines. HAN poses little to no danger of an accidental exothermic decomposition / thermal runaway, unlike Hydrazines. That means switching to HAN from MMH provides one of those rare all-around performance improvements. NTO/HAN only provides a modest Isp increase over NTO/MMH (2% to 3%), but HAN is far less toxic than Hydrazine, as it's only mildly acidic. No fuel line blow-down and double check valve is required using HAN. NTO (N2O4) is very toxic and very oxidizing / corrosive, but not quite as toxic to humans as Hydrazine-derivative fuels. LOX is also pretty nasty stuff to work with, just in case someone thinks pure O2 is significantly less problematic than NTO.

All things considered, we should probably make the switch from MMH as our fuel of choice to HAN because it's an improvement over MMH in every sense of the word, but keep the NTO (N2O4) oxidizer, despite its toxicity.

You can read more about HAN monopropellant here:
GPIM (Green Propellant Infusion Mission)

tahanson43206 · 2025-05-15 10:14:14

For kbd512 re #266

Thanks for this additional coverage of alternative monopropellants.

I changed the new topic to be more inclusive, in hopes NewMars members might want to collect information there.

I learned recently that at least one of our members does not know how to use the search tool.

(th)

tahanson43206 · 2025-05-15 14:51:35

For kbd512 ...

Picking up on your interesting ammonia idea...

Per Google, ammonia freezes at -108 Fahrenheit (-77.8 Celsius) .... that seems like a modest temperature to hold for years in space with shade.

Would that make a difference in the mass of tanks for shipment?

We have no topics about using ammonia for anything other than your topic about making ammonia as a convenient energy storage system on Earth.

I think you may have come up with an idea that is novel (in the sense that no one else appears to have written about it)

I asked Gemini and will post the answer in a new topic.

(th)

NewMarsMember · 2025-05-19 06:47:07

For kbd512 re Falcon 9 first stage improvements:
https://newmars.com/forums/viewtopic.ph … 69#p231769

This set of calculations might be interpreted as providing more payload to LEO or as requiring less propellant for the same payload.

Either way, the savings to the customer should be measurable in dollars.

The cost of the improvements will be substantial, due to the need for extensive testing to make sure all the new components hold up to service life extending to hundreds of flights. That cost will be amortized over those hundreds of flights.

It might be possible to calculate the break even point.

After the investment has been paid back, the savings might be shared with the customer.

It seems to me this must be the accounting that goes into any similar improvements in a capitalist enterprise.

(th)

Last edited by NewMarsMember (2025-05-19 06:48:02)

kbd512 · 2025-05-19 08:33:17

tahanson43206,

My proposal WILL produce more aerodynamic drag, especially in the lower atmosphere. There can be no doubt about that. The question I have is, if we deleted over 15,000kg in mass from the stage by using a mass and volume optimized stage, and improve the engines to use a staged combustion vs gas generator cycle, do we still gain additional payload performance by making the Falcon's booster stage drastically lighter?

Aero drag matters a lot to airliners that continuously operate at high subsonic speeds lower in the atmosphere.

Does it matter nearly so much to rockets / rocket boosters that are only briefly within the lower atmosphere?

For example, the real world Falcon 9 booster attains an altitude of around 60km / 196,850ft at burnout, and it only burns for 162 seconds. There's almost no atmosphere up there, and average ascent rate is 72,907ft per minute. The vehicle is traveling slower in the lower atmosphere due to TWR, then much faster as its altitude increases and propellant supply is depleted.

When the booster comes back through the lower atmosphere, my contention is that more aero drag is better, because less propellant is required to slow the booster for landing.

Do we end up gaining back some of what we lost to drag during the ascent, in terms of reduced propellant consumption for the landing burns?

Maybe we do or maybe we don't, but a 15,000kg dry mass reduction using optimized stage geometry is a very significant amount of mass to remove from the Falcon 9 rocket's dry mass. The landing gear mass would also be significantly reduced because the landing gear track is so much wider, relative to the height of the booster. There's very little risk of toppling over. What might otherwise be a failure at a landing attempt with a 3.66m diameter / 42.6m height booster core would be perfectly acceptable with a 10m diameter / 11m height booster core. The required stroke length of the landing gear would be reduced with that much mass removed. The grid fins should become proportionally more effective at maneuvering the core as well. The cost of spinning fiber around a mold is much less than the cost of machining Al-2195 gores, precisely bending them, and welding them. Recall how long that process takes ULA when Tory Bruno gave a guided tour of ULA's rocket factory, because ULA is doing the exact same thing for the Atlas V core stage.

Composite cryotank development was intended to reduce tank mass, fabrication cost, and fabrication timeline over Al-2195, and all program objectives were met. I would think that the "less extreme" LOX and RP1 propellants, would permit further mass reductions since it doesn't have to survive LH2 internal pressurization and Hydrogen permeation.

Peak heating of the Falcon 9 booster's sidewalls is up to 20kW/m^2 (12.9W/in^2), and peak temperature is 250C (482F), caused not by aero heating but impingement of the rocket exhaust during the retro-propulsive burn. That said, the composites surrounding the F-35's engine are continuously subjected to 1,000F exhaust bypass from its F-135 engine, with peaks to 1,100F. Since those things don't fail, we should not expect a composite propellant tank to fail, either, when appropriate resins are used. However, the F-35's high temp skin panels also have their high temp resin cured in a high temp autoclave, which removes porosity from the composite and improves strength and weight. This would significantly add to the cost of the booster, as compared to NASA's non-autoclave process, but it might be worth the cost because of how much durability and temp resistance is increased. SpaceX uses Falcon 9 / Falcon Heavy as a high op-tempo workhorse rocket, much like the US military uses its F-35s. If the airframe / propellant tanks are now good for a thousand flights, then you may only need a handful of them to continue providing launch services to the US government.

The bottom line is that there are tradeoffs here. I'm very curious to know if that whopping stage mass reduction is "worth it" in terms of payload performance and other secondary effects such as landing propellant reduction, reduction of landing gear mass, landing stability on the deck of a heaving drone ship, etc.

tahanson43206 · 2025-05-19 11:04:47

for kbd512 re #270

I asked Gemini if it could find OpenFOAM models that would feature cone shaped vehicle body design, but the nearest it could find is studies for nose cones. The ship design you've described is one giant nose cone, so perhaps the studies for nose cones can be adapted.

This is just a preliminary look, using the public facing Gemini. There is an advanced version that is available (occasionally) so that might be worth trying. However, OpenFOAM can certainly be set up to run a new study. I'm not ** there ** yet, but I can see that capability on the horizon. If you have a bit of time and want to search for CFD studies of nose cone performance, there might be something we could adapt.

Conversation with Gemini
Hi Gemini! Thanks for you help with previous requests! This one should be right in your ballpark, because it just involves search. Please look for OpenFOAM models for cone shaped space vessels such as DCX. The tradeoffs under study are drag at liftoff vs benefits of mass savings due to squat tank design.
***
Hello! It's great to hear from you again. I can certainly look into OpenFOAM models related to the drag characteristics of cone-shaped space vessels like the DC-X, especially concerning the trade-offs between liftoff drag and the mass savings from a squat tank design.
Certainly! Based on the search results, here's what I found regarding OpenFOAM models for cone-shaped space vessels and the trade-offs you mentioned:
OpenFOAM for Drag Analysis of Cone-Shaped Spacecraft
Nose Cone Design is Critical: The shape of the nose cone significantly impacts aerodynamic drag, especially during ascent at subsonic speeds. Different shapes like conical, elliptical, parabolic, ogive, and Von Karman are studied to minimize drag and optimize flow characteristics.
CFD as an Analysis Tool: Computational Fluid Dynamics (CFD) software, including OpenFOAM, is used extensively to analyze the drag forces acting on various nose cone designs at different Mach numbers.
OpenFOAM Capabilities: OpenFOAM is capable of simulating compressible flows and has been used for external aerodynamic applications, including analyzing flow around cones at various angles of attack. Solvers like rhoCentralFoam are used for such simulations.
Turbulence Modeling: OpenFOAM offers various turbulence models (RANS, LES) suitable for analyzing complex flow phenomena around spacecraft, including separated flows which can occur in the subsonic regime.
Trade-offs: Liftoff Drag vs. Mass Savings (Squat Tank Design)
While the search results didn't specifically address "squat tank design" in the context of cone-shaped spacecraft, we can infer some general principles and potential areas of study using OpenFOAM:
Squat Tank Implications: A squat tank design, being shorter and wider than a traditional cylindrical tank, might offer mass savings due to potentially simpler construction or better structural efficiency for a given volume. However, the wider base could lead to increased drag during liftoff, especially if the overall vehicle shape transitions less smoothly from the wider base to the upper sections.
CFD for Trade-off Analysis: OpenFOAM could be used to model the aerodynamic flow around a cone-shaped spacecraft with a squat tank. Simulations could be run at liftoff speeds to quantify the drag generated by the wider base compared to a similar volume with a more elongated tank.
Parametric Studies: By creating different CAD models with varying degrees of "squatness" for the fuel tank while maintaining the overall volume and upper conical shape, parametric CFD studies can be performed in OpenFOAM. This would allow engineers to directly compare the drag coefficients at liftoff for each design.
Optimization Potential: OpenFOAM, coupled with optimization libraries, could potentially be used to find a tank shape that balances the need for mass savings (presumably linked to a squatter design) with the requirement for minimized drag during the initial ascent phase.
In summary: OpenFOAM is a suitable tool for analyzing the aerodynamic drag of cone-shaped space vessels, and it could be effectively employed to study the trade-offs between the liftoff drag implications of a squat tank design and the potential mass savings it offers. This would likely involve creating specific geometric models incorporating the squat tank concept and running detailed CFD simulations at liftoff conditions.
***
that's disappointing. I was hoping there might be a model we could run immediately, instead of having to go through the laborious process of creating a model. Please try again! Let's see if different search terms help to uncover existing models or tutorials.
***
You're right, my apologies! I got ahead of myself in describing the capabilities of OpenFOAM rather than finding readily available resources. Let's try a more focused search for existing models or tutorials that might be closer to what you need.
Unfortunately, even with more specific search terms like "OpenFOAM tutorial cone shaped space vessel," "OpenFOAM models cone shaped space vessel," and "OpenFOAM CFD simulation cone with squat tank," I haven't found any immediately downloadable OpenFOAM models or step-by-step tutorials that directly address the DC-X or a cone-shaped vessel with a squat tank configuration for drag analysis at liftoff.
Here's what the search results do indicate, which might be helpful in the longer run:
Focus on Nose Cone Aerodynamics: There are studies using CFD (likely including OpenFOAM in some cases, though not explicitly stated in the snippets) to analyze the aerodynamic performance of different nose cone shapes for rockets and spacecraft. This confirms that the general shape is a relevant area of study.
DSMC for Spacecraft Re-entry: One result mentions using OpenFOAM's dsmcFoam solver for simulating rarefied gas flow around a space capsule during re-entry. While this is a different flight regime than liftoff, it demonstrates OpenFOAM's applicability to spacecraft aerodynamics. The geometry described is a blunt body with a truncated cone shape, which is somewhat related.
SnappyHexMesh for Complex Geometries: OpenFOAM's meshing tool, snappyHexMesh, has the capability to handle complex geometries, including analytical shapes like cones. This is relevant for creating the mesh for your specific vessel design.
Mixing Tank Simulations: Some tutorials and discussions focus on simulating flow within tanks using OpenFOAM, including the use of dynamic meshes for moving parts. While the application is different, the techniques for setting up tank geometries and analyzing fluid flow might offer some transferable knowledge for the squat tank aspect.
Lack of Direct Matches: The absence of direct hits suggests that a ready-to-run model for your specific scenario (cone-shaped liftoff drag with a squat tank) is likely not publicly available.
Where to go from here:
Given the search results, it seems the most effective path forward will involve creating your own model in OpenFOAM. While this is the "laborious process" we were hoping to avoid, the tools and general methodologies within OpenFOAM are suitable for this type of analysis.
Here are some steps you might consider:
CAD Model: Create a 3D CAD model of your cone-shaped vessel incorporating the squat tank design.
Meshing: Use OpenFOAM's meshing tools (like snappyHexMesh or blockMesh depending on the complexity) to generate a suitable computational mesh around your geometry.
Case Setup: Define the physics of the problem within OpenFOAM, including:
Solver: Choose a compressible flow solver appropriate for subsonic liftoff conditions (e.g., rhoCentralFoam).
Turbulence Model: Select a turbulence model suitable for external aerodynamics at these speeds (e.g., k-epsilon, k-omega SST).
Boundary Conditions: Define appropriate inlet (atmospheric conditions), outlet, and wall (spacecraft surface) boundary conditions.
Material Properties: Set the properties of air.
Simulation: Run the OpenFOAM simulation.
Post-processing: Use ParaView (OpenFOAM's native post-processing tool) to visualize the flow field and extract relevant data like drag forces.
Parametric Study: To analyze the trade-offs, you would need to create and simulate multiple geometries with varying squat tank dimensions while keeping the overall volume constant, then compare the drag results.
While not the immediate solution we hoped for, this approach will provide you with the specific data you need for your analysis. There are numerous OpenFOAM tutorials available online that cover the individual steps of this process, even if they don't address your exact geometry.
I'm here to help if you have questions about any of these steps as you move forward!
***
Thank you for doing the research and for preparing a summary of the results. I'll pass them along to the client.
***
You're very welcome! I'm glad I could at least provide a summary of what's out there and a potential path forward. Please let me know if your client has any further questions or if you encounter any specific challenges as you begin the modeling process. I'm here to assist in any way I can.

(th)

kbd512 · 2025-05-19 12:48:58

I found these gravity and drag losses posted here:
Drag: Loss in Ascent, Gain in Descent, and What It Means for Scalability

Ariane A-44L: Gravity Loss: 1576 m/s Drag Loss: 135 m/s
Atlas I: Gravity Loss: 1395 m/s Drag Loss: 110 m/s
Delta 7925: Gravity Loss: 1150 m/s Drag Loss: 136 m/s
Shuttle: Gravity Loss: 1222 m/s Drag Loss: 107 m/s
Saturn V: Gravity Loss: 1534 m/s Drag Loss: 40 m/s (!!)
Titan IV/Centaur: Gravity Loss: 1442 m/s Drag Loss: 156 m/s

As you can see, gravity losses tend to dominate.

The Saturn V has a 10m diameter core, but it also has a very different length-to-diameter ratio.

This blurb from that "GravityLoss.com" blog post seems to lend some credibility to my assumption that a "stubbier" rocket would require less descent propellant:

Meaning that the terminal velocity is inversely proportional to the square root of the length of the rocket l and the density of the almost empty rocket, rho_ter. This means that if density is held constant (a reasonable first assumption), as well as shape (similar C_d), the terminal velocity is doubled when the rocket size (length) is quadrupled. And this means more delta vee needed for landing. It doesn’t matter if the rocket is lengthened and widened or only lengthened. If the constant density is held, the increased length (or depth) just means more mass per frontal area.

Edit:
The website seems to be run by a fellow space enthusiast named Valtteri Maja. The data / analysis comes from Jon Goff (co-founder of Masten Space Systems, President and CEO of Altius Space Machines), from a blog post over on Selenian Boondocks, which is another space / Thorium power enthusiast website run by former NASA employee Kirk Sorensen. That means it has some credibility behind it.

Orbital Access Methodologies Part I: Air Launched SSTO

Last edited by kbd512 (2025-05-19 12:57:38)

tahanson43206 · 2025-05-21 11:39:02

For kbd512 re On-Orbit defense system...

Impressive image!

My understanding is that China has expressed concern.

One interpretation might be that this is a good sign!

Your new topic should have plenty of running room, over coming years. This is a marathon and not a sprint.

(th)

tahanson43206 · 2025-06-01 20:54:38

For kbd512 re your new ITV topic ...

Here is an attempt to try to capture at least ** some ** elements of your idea.

(th)

tahanson43206 · Today 07:53:55

For kbd512 re post in ITV about carbon structure...
https://newmars.com/forums/viewtopic.ph … 72#p232072

I'm wondering if an oil derrick might be a model for the ITV?

GW is interested in the forces that would tear the swing arm from the rocket fuselage, and you've proposed carbon as a building material for the vessel.

An oil derrick is a system of straight segments that is able to spread force over a wide base compared to the narrow peak.

Such a structure, if done with carbon, might have the resilience to withstand forces that would rip a metal structure.

This forum could use a member with the ability to create an architectural plan, and we are open to applications for membership.

The ITV project will remain in a preliminary phase until we have someone who can turn words into plans.

A plan would be generated with one of the many CADCAM tools that exist, and renders would allow NewMars readers to see what the design would look like if assembled in orbit.

(th)

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