Discussion Orbital Propellant Tanker

T.Neo

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OPT-MKII

Using some revised mathamatics, I present the Orbital Propellant Tanker Mark II, with revised dimensioning and some more carefully calculated mass figures:

attachment.php


OPT-MKII stands almost 77 meters tall and is roughly 10.5 meters in diameter.

Changes include:

- Widening of tank diameter to 10.5 meters up from 8.4 meters (the result of an attempt to have pointless commonality with the Shuttle ET). A small increase in width can result in a large decrease in height- improving the ease of vehicle integration, transport to the launch site, and launch pad operations.

- Mass figures are more carefully calculated, including more attention paid to insulation foam.

- Extra engine added (up to six from five) to increase liftoff thrust, decrease gravity losses during the early portion of the ascent, and add non-RCS roll-control later in the ascent.

However, the main difference here is the inclusion of a 1.5 staging system as seen on the Atlas rocket. This makes OPT conceptually similar to the National Launch System, a shuttle-derived concept studied in the '80s by NASA and the USAF. The NLS-2 was a vehicle derived from an STS ET, powered by 6 hydrolox 'STME' engines, derived from the SSME.

nls2m.jpg


Also of note is the Saturn V-B, a 1.5 stage launcher made up of an S-IC derived stage (the S-ID), which could lift over 22 tons to LEO. The S-ID was also intended for use in other, multi-stage Saturn studies to increase payload, and is simulated in the Velcro Saturns addon.

Why drop four engines? Well, the answer is simple: after a certain amount of propellant has been burned, they are simply not needed. A five-engine OPT reaching burnout would experience very high accelerations, perhaps in the range of 10G or more, even with all engines throttled down to the lowest possible setting.

The solution then, to reduce stresses on the vehicle's lightweight structure, would be to shut down some engines. However, these engines would then become dead weight and do nothing but serve to reduce capability unecessarily.

A single RS-68 engine masses 6.6 tons, and there would be multiple uneeded engines. In addition, some of the piping and structure needed to support those engines would also be dead weight.

Ergo, the 1.5 stager is born. Extra mass is added in the seperation system for the discarded 'engine pack', but the mass of this system that remains on the vehicle after seperation is far less than the mass of four engines and their support structure. Disconnecting propellant feed lines are nothing to fear- they have been practiced for 30 years with the Shuttle.

An extra engine can be added. Although this engine will increase the amount of mass the vehicle needs to take to orbit, it will pay for itself by improving performance early in the flight and reducing gravity losses. In addition, it will allow a degree of roll-control after seperation, negating the need for a seperate thruster system to do the same job.

After seperation, the two remaining engines can be throttled down as required to achieve optimum acceleration.

A 1.5 stage vehicle offers improvments over a vehicle using parallel-burning boosters, for example, by offering increased production runs of common components (namely engines), therefore reducing cost and increasing reliability.

Depending on seperation velocity and engine properties, it might be possible to recover and reuse the engine pack. The booster pack of a 1.5 stage vehicle being mainly engines and their support structure, could be easier to recover than a flimsy mass-optimised propellant stage.

The performance gain by seperating unecessary engines can be used to:

1. Increase payload.
2. Ensure a viable mass ratio or make the vehicle structure sturdier.
3. Reduce vehicle size.
4. Two or more of the above.

The common propellant tank, however, is justified by easier handling, construction, and integration, less "wasted mass" in the late portions of ascent, and potential on-orbit applications.

A whole family of vehicles is potentially possible by stretching or "squashing" the propellant tanks. There is no reason not to include conventional payload-launching vehicles in this potential family, as well as derivatives of the Orbital Propellant Tanker concept itself.

The propellant tank is still white. This is in an attempt to slow propellant boiloff on orbit. A payload-launching design would leave the tank unpainted, and a signature orange colour (or grey... or light green... it depends on the insulation used).

Another new addition is the nosecap, which is not actually intrinsic to the MKII design. It is larger, but aerodynamically far superior to the silly nosecap of the first version (see the beginning of this thread).

Comments, suggestions and criticism are welcome. :cheers:
 
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T.Neo

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Aft section and breakaway engine ring

Theoretical work on the aft section continues. I've roughly worked out... very... simple and ugly propellant line layouts, but not the actual engine support structure itself. What makes up the engine support structures of most other space vehicles? Is it an internal framework covered by thin metal sheeting? Or some sort of monocoque construction? Is it usually constructed out of aluminium or steel? What material is used for propellant piping? How do these factors differ among different vehicles and what are their pros and cons?

attachment.php


I've tried to take care to ensure enough space for the center engines during seperation, also it's obvious that the engine ring would be hit by the exhaust from the center engines... would obviously need some sort of TPS if it was intended to be reusable.

I haven't gotten to the other stuff that has to fit in there yet, and I haven't thought about where the fill/drain lines would have to be situated...

Nothing is textured yet. I'm quite a ways away from the graphics side of things, that is why for example the details of the RS-68 engine are obscured by big red boxes. They're just there to see if the rough dimensions of the engines would conflict with any other stuff in that internal volume.

This is also an interesting image of the STS-1 ET after ascent, showing clearly the marks from exhaust heating the aft section during ascent, as well as from the SRB sep motors and aerodynamic heating:

s81-30509.jpg


That might give a good idea of what a vehicle like this might look like after the rigours of ascent...
 
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n72.75

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Food for thought:

Imagine the spacecraft analog:

rtc26_delivery.jpg
 

T.Neo

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Awesome! but the textures make the rocket look more of a toy rocket than a feminine hygiene product.

Oh heavens no, I would never make textures like that... those are just different materials, so I can distinguish the different components. :lol:

500px-Cad_mouse_1.svg.png


Nobody would make such a garish mouse, either. :p
 

T.Neo

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attachment.php


Are my engines spaced too closely? They look roughly as closely-spaced as the engines on the NLS, and nozzles on other launchers are also quite closely spaced, but Ares V apparently had engine spacing problems as described in this article. The resulting arrangement had quite widely spaced engines:

5481_single.jpg


index.php


I can arrange my engines differently than I have them now, but this will affect some of the internal geometry.

I'm also not sure if my nozzles are the right size- I'm going by the 96 inch diameter figure stated on the Pratt & Whitney site, presumably this refers to the end of the nozzle bell (I can't really think of anywhere else that could be described by "diameter").
 
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OrbitalPropellantTanker.jpg

An idea I had a week or two ago, a launch vehicle exclusively for delivering liquid hydrogen propellant to orbit, to supply a fusion spacecraft or similar. This concept uses hydrolox propulsion with a common bulkhead between the LOX and LH2 tanks; the payload is contained in the same tank as the fuel.
Not sure why it is called "orbital", as it isn't an orbit-only vehicle, it doesn't go from orbit to orbit or anything like that; effectively it's just a species of launch vehicle. But it does launch to orbit, I guess.
The sunshade isn't shown or described in the document, but I've envisioned it as basically a sheet of solar sail material (not for any propulsive purpose, it's just lightweight and reflective) extended on spring-loaded booms, somewhat similar to long, thin antennas. I know this arrangement doesn't provide 100% effective shielding, but for what it's worth, it provides at least some shielding.
Don't take all of the numbers too seriously, a lot of the smaller mass breakdown figures are more or less thumbsucks by me, and the cost per kg is at best very rudimentary. Nevertheless, if you spot something that's blatantly, obviously, horribly incorrect, please tell me.
Overall I generally need a cheaper cost/kg figure, which would probably require a less... mundane system, but for what it's worth I think it is rather nice. I'm sure it has a use somewhere in the Orbitersphere.
The only real downside, is that it ended up looking like a feminine hygiene product. :facepalm:


This can work. But I think the numbers you are using for the structural components are too optimistic. For instance the value you are assigning for tankage for the mass of propellant is too small.
See the equations here for more realistic numbers:

Mass Estimating Relations
• Review of iterative design approach
• Mass Estimating Relations (MERs)
• Sample vehicle design analysis
http://spacecraft.ssl.umd.edu/academics/483F09/483F09L13.mass_est/483F09L13.MER.pdf


Bob Clark
 

T.Neo

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I am fully aware that the mass figures in the .pdf posted at the beginning of this thread are wildly unrealistic. The whole point of the MKII is to use plausible mass estimates and I am investing a lot of effort and time into insuring that this is so.

Currently I am having more issues with engine support structures, piping and subsystems than tankage.
 

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It should be able to tether itself to the ISS to serve as a refueling basis for inter planetary missions.
 

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But it does launch to orbit, I guess.

1341154/109154 = 12.287

Ve in vacuum is 4.120 km/s, Ve at sea level is 3.580.65 km/s.

So the delta V will be somewhere between
ln(12.287) * 4.12 km/s and ln(12.287) * 3.58065 km/s

At first glance delta V between 10.34 and 8.98 km/s seems more than adequate to achieve LEO.

While RS-68A has a nice ISP, thrust is weak. At sea level 5 RS-68a engines would impart 15700000 newtons. Since mass is 1341154 kg, acceleration is 11.7 ms-2. It is desirable to climb above the thick atmosphere before achieving orbital velocity. So most launches start with a vertical climb.

Gravity's acceleration is about 9.8 ms-2. Each 102 seconds of vertical ascent levies a 1 km/s penalty.

You start your vertical ascent with a net acceleration of (11.7 - 9.8) km/sec^2. Or 1.9 km/sec^2. With this very slow ascent, gravity losses between 2.5 and 3 km/s are likely.

Gravity loss breaks your design. It's unlikely your rocket would achieve orbit.

----

Also your liftoff mass/empty mass ratio is 22.7. Even if you had dense kerosene, I'd regard that as difficult for a single stage vehicle. When you must contain 3363 cubic meters of super cryogenic hydrogen, it is even more implausible.
 
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T.Neo

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HopDavid: I suggest you stop complaining about your prejudice against the RS-68, and start complaining about the fact that the vehicle described in the .pdf at the beginning of the thread, has a propellant tank weighing only 3 tons. :tiphat:

I really do not know what you are complaining about, because I have made a simple version of that vehicle and flown it to orbit twice in Orbiter (albeit using a two-burn trajectory, but this was more due to the vehicle's odd acceleration curve and my bad piloting skills).

Also, Jon Schilling's Launch Vehicle Performance Calculator states that the vehicle would reach orbit with a considerable margin.

I would not call the RS-68 "anemic". It certainly has enough thrust for the job, and a higher ISP than competing kerolox engines. Adding another RS-68 would improve thrust and therefore gravity losses.

Yes, with 5 RS-68 engines, that vehicle is a slow riser. But this does not stop a vehicle from working. The Saturn V was a pretty epic slow-riser as well. It worked just fine.

If there is an error that makes that vehicle incapable of reaching orbit, it is likely pretty minor, rather than glaringly obvious.

Also your statement of "super cryogenic hydrogen" makes me doubt your knowledge of where these engineering difficulties actually lie. Yes, LH2 is cold and yes, that poses an engineering challenge. But as far as I can understand, the greater impact on tank weight is the low density of LH2.

So please do not point a finger and shout "you do not understand the problems of SSTO!" and then talk about gravity losses. High gravity losses are not intrinsic to SSTOs.

Nevertheless, it is completely pointless. The MKII vehicle is totally different from the first concept, making pretty much every criticism of the MKI null and void.

The MKII has six RS-68 engines. And no, it doesn't have a 3 ton propellant tank holding 1200 tons of propellant. ;)
 
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HopDavid

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Yes, with 5 RS-68 engines, that vehicle is a slow riser. But this does not stop a vehicle from working. The Saturn V was a pretty epic slow-riser as well. It worked just fine.

This graphic comes from a NASA page on Apollo history:
gforce.gif

Scroll down towards the end and it's right after T = 000:11:39.

After 160 seconds, acceleration has reached about 38m/sec^2. In contrast, after 160 seconds, your vehicle will have reached an acceleration of around 26 m/sec^2.

Which is to be expected, the 1st stage of Apollo was lox/kerosene.

The upper two stages (which don't need to fight as much gravity loss) were lox/hydrogen.

So please don't cite a three stage vehicle with the lower stage using kerosene to defend your single stage to orbit hydrogen vehicle.

Also your statement of "super cryogenic hydrogen" makes me doubt your knowledge of where these engineering difficulties actually lie. Yes, LH2 is cold and yes, that poses an engineering challenge. But as far as I can understand, the greater impact on tank weight is the low density of LH2.

Oops. You snipped some context. What I actually said was "When you must contain 3363 cubic meters of super cryogenic hydrogen"

That huge volume comes from the low density of hydrogen.

So please don't employ your sloppy reading skills to claim I ignored that failure of your design. Such a huge flaw was hard to miss.

One thing I did ignore though, the huge volume of hydrogen will likely increase your cross sectional area and thus increase atmospheric drag.

So please do not point a finger and shout "you do not understand the problems of SSTO!" and then talk about gravity losses. High gravity losses are not intrinsic to SSTOs.

Low T/W ratio and high gravity losses are intrinsic to high ISP engines.

For example the T/W of the RD-180 kerosene engine is 77.26

RL-10A-4-2 Lox/hydrogen engine is 60.53

Astronautix doesn't list the RS-68a but they do list the T/W for the RS-68 is 51.2

Ion engines have even better ISP. These can weigh hundreds of newtons but have a thrust less than 1 newton. With this very poor T/W, an ion engine couldn't even get off the ground. If you needed a vertical ascent from a massive body like the earth, gravity loss would consume all of an ion engine's propellant.

So if you want an SSTO with high ISP, you're going to suffer gravity loss.

If you want an SSTO with good thrust, you're going to suffer ISP loss.

Using lox/hydrogen engines doesn't make SSTO magically easy.
 

n72.75

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Enough.

If we can't have a reasonable debate about an engineering topic without criticizing each others intelligence and reading skills, we shouldn't even have the discussion.

@HopDavid, If you have a good idea for a propellant tanker, why don't you produce an add-on. Then third parties can decide which is "better".

I for one have toyed with the idea of developing a tanker for several months and find many of the ideas here thought provoking and interesting.
 

T.Neo

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I'm just using an example of the Saturn V as a vehicle with high gravity losses. Since it and an SSTO have to ascend in pretty much the same way, it does not really matter that the high gravity losses are at the beginning of the first-stage ascent.

HopDavid, let's get down to the crux of the matter here:

There are two options for propulsion here (there are others, but they are ignored out of practicality in this case). They are hydrolox and kerolox. Hydrolox has the better ISP, but kerolox generally has better thrust and you complain about the thrust/weight of hydrolox engines.

If you think one or the other (in this case the alternative, which is kerolox) is better, then please make your case as to why it would be physically, economically, and technologically superior for application on a vehicle like this.

I have actually considered a kerolox-propelled vehicle (which would make this similar in concept to the Saturn V-B). As well as a vehicle utilising two types of engine, burning two different propellant combinations.

And there are also things like Thrust Augmented Nozzles. TANs on a vehicle such as this could both boost payload, and use propellant tanks jettisoned with the breakaway booster pod... would improve performance nicely.
 

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There are two options for propulsion here (there are others, but they are ignored out of practicality in this case). They are hydrolox and kerolox. Hydrolox has the better ISP, but kerolox generally has better thrust and you complain about the thrust/weight of hydrolox engines.

If you think one or the other (in this case the alternative, which is kerolox) is better, then please make your case as to why it would be physically, economically, and technologically superior for application on a vehicle like this.

For vertical ascent to upper atmosphere, kerosene's better. For achieving orbital (or escape velocity) above earth's atmosphere, hydrolox is better.

Unfortunately with single stage it's hard to have both. Having two types of propellant and two types of rocket engines in the same stage would increase complexity.

That is why I favor two or three stages to orbit. ULA's Atlas/Centaur is a good combo in my opinion. It has launched more than a few com sats. The ULA folks mention their extensive experience with hydrogen rockets: the Deltas and Centaurs. The ACES upper stages would incorporate aspects of both.

For propellant depots, ULA's ACES architectures are the best researched and most well thought out, in my opinion.

Here are some links on ULA's proposed depots:
Depot Based Transporation Architecture
A commercially based lunar architecture

In the Commercially Based Lunar Architecture, they mention other players besides ULA could deliver propellant to the depots. It would be good to see SpaceX Falcon rockets delivering propellant. Or even suppliers as well as depot users from other countries. With multiple suppliers and consumers, orbital propellant would move closer to being a competitive commodities market.
 
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T.Neo

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Three stages to orbit increases seperation events and complexity. Two stage to orbit is optimal, but not required.

The arrangement I have here allows common use of a single engine type, single propellant tank, and removes all air-start events.

I have news for you: gravity losses are not the main problem of SSTOs. And this is also not an SSTO. It does not need to be. It makes no sense for this vehicle to be an SSTO. There's some 50 tons of mass that halfway to orbit you just don't need anymore. So you throw it away.

I can stage and not need a wholly different upper stage with wholly different engines and tanking.

Also, the ACES depot stuff is nice... but not intended to fulfill the same role as this thing. Which was originally to supply propellant to my absurd fusion-powered interplanetary vehicles- a task naturally requiring a very high upmass capability.

EDIT:
(Also: you have not seen the 2.5 stager derivatives of this thing that I've also been drawing up.)
 
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HopDavid

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I have news for you: gravity losses are not the main problem of SSTOs.

Gravity loss would be a problem if your SSTO uses hydrolox.

Another problem is achieving the achieving the tiny dry mass/fully loaded mass fraction needed to achieve a delta V budget. It wouldn't be possible to enclose 3363 cubic meters of hydrogen with the mass you allocate.

When delta V budgets require impossibly small mass fractions, this can be circumvented by throwing away mass during the trip in the form of expendable rocket stages.

And this is also not an SSTO. It does not need to be. It makes no sense for this vehicle to be an SSTO. There's some 50 tons of mass that halfway to orbit you just don't need anymore. So you throw it away.

I fully agree. So why did you title your pdf "Single-Stage-To-Orbit non-reusable propellant tanker."?

Also, the ACES depot stuff is nice... but not intended to fulfill the same role as this thing. Which was originally to supply propellant to my absurd fusion-powered interplanetary vehicles- a task naturally requiring a very high upmass capability.

You'd have more upmass capability if you discarded unneeded deadweight along the way. Should we ever use hydrogen as reaction mass in non chemical rockets, I believe the ACES tankers could supply these rockets as well.

EDIT:
(Also: you have not seen the 2.5 stager derivatives of this thing that I've also been drawing up.)

I expect I would have much less problems with these.
 
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T.Neo

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Gravity loss would be a problem if your SSTO uses hydrolox.

The relationship between engine T/W and vehicle T/W is what you must look out for. Mass of the engines pails in comparison to the mass of the propellant, so having more massive engines does not really matter if it affords you extra thrust.

Engine mass is more critical in terms of directly stealing away from payload, but this is a rocket equation issue, not a T/W one.

I have not crunched the numbers, but I overwhelmingly suspect that a hydrolox SSTO will outperform a kerolox one. It's a matter of ISP.

Another problem is achieving the achieving the tiny dry mass/fully loaded mass fraction needed to achieve a delta V budget. It wouldn't be possible to enclose 3363 cubic meters of hydrogen with the mass you allocate.

Have I not repeatedly made fun of the 3 ton propellant tank? ;)

When delta V budgets require impossibly small mass fractions, this can be circumvented by throwing away mass during the trip in the form of expendable rocket stages.

Not necessarily expendable and not necessarily entire stages.

See Atlas, the original originator of the concept I am using here.

I fully agree. So why did you title your pdf "Single-Stage-To-Orbit non-reusable propellant tanker."?

Because the vehicle described in that PDF was an SSTO (and also requiring magical propellant tanks).

This vehicle is no longer an SSTO.

You'd have more upmass capability if you discarded unneeded deadweight along the way.

ACES/Atlas does this and has lower upmass. I'm not complaining about upmass as a whole, but the upmass of Atlas particularly. This is not Atlas.

And multiple factors can effect cost. When you are talking about many hundreds of tons of propellant being used each year, it is pretty helpful to reduce costs.

Should we ever use hydrogen as reaction mass in non chemical rockets, I believe the ACES tankers could supply these rockets as well.

ACES is awesome. This is not ACES.

Why don't you make an ACES addon?

I expect I would have much less problems with these.

Because their designs are less unusual? ;)

The upper stages I have in mind are relatively small 'kick stages' that only boost payload by about 20-30 tons.
 

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Here's a proposal for low cost hydrolox SSTO's to deliver propellant and other expendables to LEO:

Aquarius.
http://www.astronautix.com/lvs/aquarius.htm

The developers intent was to use low cost pressure-fed rockets that would not have the same reliability level as current rockets. But they would only deliver expendables so the expected number of launch failures would just be made up for with further launches.
I do have a problem with their proposed design. On that Astronautix page you can see for their propellant mass it's 117.5 tons and their propellant tank mass it's 3.7 tons. This is a tankage ratio in the range of 30 to 1. The problem is this a tankage ratio that would be expected for pump-fed engines which have much lower tank pressures than pressure-fed engines.
In fact their design would use pressures in the tanks up to 30 bar. This is 10 times or more greater than the pressures in pump-fed tanks. So you would expect the tank mass to be 10 greater than the amount they are using.
They do say they are using composite materials for the tanks, but even optimistically this would reduce the tank mass by half. So you are still at 5 times the tank mass they are using.
However, if you used pump-fed engines, you would improve your Isp as well as bring the tank mass in the realm of what they are using.


Bob Clark
 
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Here's a proposal for low cost hydrolox SSTO's to deliver propellant and other expendables to LEO:

Aquarius.
http://www.astronautix.com/lvs/aquarius.htm

The developers intent was to use low cost pressure-fed rockets that would not have the same reliability level as current rockets. But they would only deliver expendables so the expected number of launch failures would just be made up for with further launches.
I do have a problem with their proposed design. On that Astronautix page you can see for their propellant mass it's 117.5 tons and their propellant tank mass it's 3.7 tons. This is a tankage ratio in the range of 30 to 1. The problem is this a tankage ratio that would be expected for pump-fed engines which have much lower tank pressures than pressure-fed engines.
In fact their design would use pressures in the tanks up to 30 bar. This is 10 times or more greater than the pressures in pump-fed tanks. So you would expect the tank mass to be 10 greater than the amount they are using.
They do say they are using composite materials for the tanks, but even optimistically this would reduce the tank mass by half. So you are still at 5 times the tank mass they are using.
However, if you used pump-fed engines, you would improve your Isp as well as bring the tank mass in the realm of what they are using.


Bob Clark

I had forgotten that one of the developers had earlier sent me an update on their proposal that instead uses pump-fed engines. The abstract to this report on the new version of the vehicle is here:

Pressure-Fed versus Pump-Fed Propulsion Trade for the Aquarius Launch Vehicle.
Andrew E. Turner*
Space Systems/Loral, Palo Alto, CA 94303
http://pdf.aiaa.org/preview/CDReadyMJPC09_1980/PV2009_4898.pdf

In the full report, they acknowledge it would be a technical challenge to meet the goal set for tank mass using the pressure-fed design:

Pressure-Fed versus Pump-Fed Propulsion Trade for the Aquarius Launch Vehicle.
"Table 1 summarizes the large-scale changes to the vehicle generated by the transition from pressure-fed to
pump-fed propulsion, which was evaluated by Aerojet as mentioned in the Introduction. While it is common
knowledge that substitution of a pump-fed for a pressure-fed system can significantly decrease vehicle size, this
paper quantifies this reduction on the basis of the trajectory analysis results presented in the preceding section.
Propellant tank mass decreased from 2.2 to 1.8 MT, which is facilitated by the reduction of the volumes of the
tanks by nearly half due to the reduction in required propellant masses from 120 to 64 MT. The tank mass was not
reduced in proportion to the reduction in propellant mass, which enables the pressure-volume over weight (PV/W)
parameter of the propellant tanks to be reduced in turn. PV/W was estimated to be 1.8 million inches (46,000 m) for
the tanks in the pressure-fed vehicle, but is reduced to less than 200,000 inches for the pump-fed vehicle. In
addition to the reduced quantity of propellant, the tank pressure is expected to be reduced to below 5 atmospheres.
The reduction in PV/W minimizes the technology challenge in tank production and permits the use of modest
pressures to provide pressure-stabilized tanks instead of the higher pressures required to overcome combustion backpressure.
The length to width ratio of the largest tank, which stores LH2, is reduced from 5.3 to 2.6, which is a more
optimal shape for a pressure vessel (private communication, Microcosm Corp.)."
p. 3.

In the new version, they get a hydrogen-fueled SSTO reduced in size to 70 mT gross mass and 4.9 mT dry mass capable of 1 mT payload.


Bob Clark
 

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Slow and painful transition from baby-toy coloured design mess, to (hopefully) graphically acceptable addon graphic.

So far only the propellant tank has been textured, the other parts are still defined by garish materials. Still getting the hang of UV mapping and correlating an image on a flat picture to something that hopefully looks half-good on the actual mesh.

attachment.php


I have done some work on the nosecap though, before now it has just been "magical thing that pops off the top of the rocket".

Also, what colour is the inside of the RS-68 nozzle?

This picture shows it as a sort of reddish colour:
d4_rs68_12.jpg


This picture however shows the inside of the nozzle as a greyish colour with repeating whitish patterns inside (also some interesting detail about the aft end of the rocket, should be useful for texturing):
Delta4CBC_AFSMM2009RK_01.jpg


I'm also wondering how much detail of piping and such I will need to display of the engines. Maybe I need to get back to that RS-68 model I was busy with. :shifty:
 
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