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Suborbital refuelling

RanulfC

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Martin, let me start off with apologizing for taking this off on a false tangent. Re-reading the thread I realized I misinterpreted what you'd written and had taken your objections out of context and distored in my head, (and then on-line) what you'd said. My only excuse is that I was apparently much more ill than I thought and once my head cleared I now see where I took off on my side-trip. (Never post when you're running a high fever, even if you at the time don't KNOW you have a high fever...It was a fun weekend lets say :) )

Shall we move on with the discussion and try and forgive me for my mind slip :)

Randy
 

RanulfC

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Even tho I have a clear bias for anything winged over a plain rocket thing, I see some problems with that system. Archibald, just see it as my habit of "chercher la petite bête" in things that interest me, plus I don't get half of the tech things posted here, So please correct me.

I'm someone who tends to have no 'prefered' system, (though I suppose it would appear reading this thread that might not be true, or I just like to argue, take your pick :) ) but I like to shake things out if at all possible.

Wouldn't there be a big development cost and complexity disadvantage with this solution over something like SpaceX ?

Depends really, as I noted above size tends to drive costs as a major factor. In the mid-90s it was proposed the 1/2 size "Blackcolt" prototype would cost around $100 million over three years to reach Mach-10-ish where it would stage a STAR-48V with a 1,000lb satelllite to LEO. That used a lot of 'off-the-shelf' systems though so costs for a full-scale Blackhorse would be more than twice as much, likely three or four times due to engine development costs since it needed a new keroxide engine. As Martin eluded to you can get really expensive chasing the 'air-craft-like-operations' chimera so it's a balancing game for trade offs and advantages over compromieses. How much propellant you can transfer and how fast is going to be the key factor that drives everything else.

Once it's done, I can understand the advantages of being able to takeoff and land almost anywhere , but the investment needed for getting a vehicle that works both in atmosphere and in orbit is… big. I mean , that was one of the problem with the Shuttle, and even though it was just a glider, this one would also have to be able to fly by itself, so again adding complexity to a known complex and expensive to develop thing.

This would be a very high performance aircraft that can (barely) reach orbit and return than something that works well as a spacecraft. The Shuttle was a spacecraft, (and a pretty long duration one in orbital power and life support requirments) that could function as an adequate aircraft once back in the atmosphere. It's complex yes but in technical terms it's actually pretty straight forward as a design since what you need for, say, a hypersonic aircraft would work well for a suborbital/orbital spaceplane in general terms. It helps somewhat in that you don't actually have to design to the standards of a hypersonic cruise airframe but are only 'passing through' going up and down with your actual design realm being sub and supersonic flight.

It's not going to be cheap by any means but starting out with things like assuming a titanium airframe goes a long way towards addressing some of the problems.

Have to think of the landing gear, the complex aerodynamics , the flight control systems, the refueling system, the reentry …ect...
plus this time an autonomous atmospheric propulsion, and it makes a very BIG and expensive program.

Not sure what you mean by "autonomous atmospheric propulsion"? But OTHER than the refueling system you need to do about the same work for any other reusable concept so it's not qute that major of a change. Where things will get expensive is if you try and design and build to an 'aircraft' like standard as we've already seen. :)

The advantage I see in plain rockets or thing like SpaceX is that it can do with almost no airframe. it just stacks engine/fuel/payload. I know it's more complex than that, but you see what I mean. You do without the complexity and investment of designing a complicated airframe.

Well "plain rockets" are a bit less complex since they don't deal with larger aerodynamic ranges like a winged vehicle would but it greatly depends on the mission and flight profile. (Which is why Starship is going to be more complex than say the Super Heavy booster it sits on) So obviously something that goes from zero on a runway to suborbital velocity and back using a lifting trajectory and flight profile is going to be tougher than something that pretty much goes straight up and back down, (simplification) all things being equal. As the design doesn't seem to spend a huge amount of time in the high supersonic/low-hypersonic regime going up or coming down that makes things a bit easier but it won't be trivial.

If done, I think it would be for the military, first cause it would be so expensive to develop that the only $$ source could be a US gov order for the military, second they would do it for the advantage of having something easy to deploy and use in short time compare to a rocket.

Well "some" military types would see it as easier and more operationally flexible but others may not as the whole concept of VTVL is pretty 'responsive' as well. Getting the military to pay for it would be good but I think it might have civil applicability as well.

That said, seeing this thing fly and refuel would be fantastic.

Yep :)

Randy
 

RanulfC

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I am going to add my 2 cents with regards to in-flight refueling:
- Fuel moving between two vehicles has a momentum of its own. If vehicles are not docked, RCS will need to fire continuously to compensate.

The idea is to 'dock' using the refueling booms which have a wide degree of motion possible with the flexible joints though they won't be engineered to take the full loading as you want that as a backup capabiilty for an emergency 'detach' mechanism.

- What is the fastest that anybody has ever docked in free-fall? The best I can find is Apollo, about 7 minutes from CM disconnecting from the upper stage, until it docked with LM. And docking part was more than half of that IIRC.

16 seconds? (
) and keep in mind both sets of booms will be using active seeking and attachment systems. But the 'gripper' arrangment is about right though the booms would be using more flexible 'ball-joint' systems as they need to fold up to fit into the vehicles. (And it won't likely use electrostatic or magnentic 'grip' so there's that as well)

The 'docking' is likely not the hard part as getting a clean connection with the fueling lines is going to be both more complicated and simple at the same time and highly dependent on the amount of propellant that needs to be transfered.

Randy
 

martinbayer

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Martin, let me start off with apologizing for taking this off on a false tangent. Re-reading the thread I realized I misinterpreted what you'd written and had taken your objections out of context and distored in my head, (and then on-line) what you'd said. My only excuse is that I was apparently much more ill than I thought and once my head cleared I now see where I took off on my side-trip. (Never post when you're running a high fever, even if you at the time don't KNOW you have a high fever...It was a fun weekend lets say :) )

Shall we move on with the discussion and try and forgive me for my mind slip :)

Randy
Randy,

no problem - I wasn't exactly my kindest, gentlest, most patient, best self. My apologies for that. Glad to hear you're doing better :).

Martin
 
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Archibald

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Some words about costs of this thing, taken from varied sources

a - Probable tomorrows - how science and technology will transform our lives in the next twenty years. Marvin Cetron, Owen Davies


SPACECAST 2020, Volume 1 - Global reach-white papers - Spacelift: Suborbital, Earth to Orbit, and On Orbit – p.163 – 221 – 1994

«The system is nothing if not cheap. Clapp's target is to build a fleet of five to ten spacecraft, which could fly one Black Horse mission per day. The squadron's total operating budget would be around $100 million per year, roughly the same as that of the SR-71 Blackbird high altitude reconnaissance aircraft. That translates to a launch cost of only $500 per pound of payload, or $1,100 per kilogram; with a larger fleet, the price could fall from there.

One way to pare the cost still further is with an imaginative two-ship mission. One vehicle carries the cargo, the other flies with only its crew. It turns out that the deadheading spacecraft arrives in space with 20,000 pounds (9,100 kg) of unused fuel. So pump the fuel into the cargo carrier and fly the companion vehicle back to Earth. The second Black Horse can then proceed to orbit with a useful load of 12,000 pounds (nearly 5,500 kg), plus its crew. If the launch cost of one vehicle carrying its standard payload comes to $500 per pound ($1,100 per kg), the cost for the combined mission drops to less than $85 per pound ($187 per kg)! And at that price, many missions become possible.»


b - The Flock Booster Architecture – Low Cost Access to LEO via Sustained Fueling

Alan Goff, Novatia, Folsom, CA

For a mass budget of 500 kg per passenger, a 30-ton payload supports 60 paying passengers. At $50,000 a ticket, revenue per flight is $3M, and amortized launch costs of $1M yields a net income of $2M/flight. Assuming one flight/day, 5 days/week, 40 weeks/year, it would take 15 months to exhaust a market of 15,000 customers, generating enough income to pay for the production costs of the rocket plane and perhaps the passenger compartment.

A criticism of the cost model is that turning around a reusable rocket is a time-consuming manpower intensive operation, and doing dozens at once would scale badly. While true for current rockets designed for performance, Flock rocket planes can be designed for low maintenance and fast turnaround. The specific goal of achieving airline like operations is to control these costs. The estimates used in this cost model are engineering judgments based on current technology and existing systems.


c - Economics of Separated Ascent Stage Launch Vehicles

Chris Y. Taylor Jupiter Research and Development, Houston, TX 77043

Developed structure-payload mass ratio, rD, is the mass of new vehicle structure that must be developed divided by the payload mass. One way to reduce the developed structure-payload mass ratio is to reduce the overall vehicle structure-payload mass ratio, r. If a smaller launch vehicle is built, then it seems natural that it would cost less. If the developed vehicle mass for is reduced by reducing the total vehicle mass for a fixed payload mass, then certain recurring launch vehicle costs would be reduced as well. Unfortunately, reducing vehicle structure-payload mass ratio usually requires saving vehicle weight or improving vehicle performance which increases the vehicle development structure cost. For a launch vehicle design, there will be a vehicle structure-payload mass ratio that provides a minimum development cost. Cost savings resulting from a vehicle smaller than this optimum structure payload mass ratio would be more than offset by increased costs incurred in the more difficult development program.

Fortunately there are other ways to reduce the developed structure-payload mass ratio without having to reduce vehicle’s structure-payload mass ratio. Developed structure-payload mass ratio is only equal to structure-payload mass ratio if the development program must develop the entire vehicle. If not all of the launch vehicle has to be developed, then not all of the vehicle mass has to be counted as part of the developed structure-payload mass ratio.

For example, if a good existing upper stage exists then that stage might be combined with a newly developed lower stage. Only the mass of the new lower stage would need to be counted when calculating the developed structure payload mass ratio. While this “new launch vehicle” would really only be a partially new vehicle, there is no advantage to having a new upper stage for the sake of newness if the replacement of the old one cannot be justified economically. Similarly, existing solid rocket motors might be used as solid rocket boosters for an otherwise new launch system in order to reduce the size of the development program. In addition to reusing old stages to reduce the developed structure-payload mass ratio, smaller vehicle components can be reused as well. Engines, structure, electronics, or other parts that have been proven to have good performance in past launch vehicles might be reused in future ones with the development program focusing only on portions of the vehicle where the cost of a new component can be justified by its increased performance. A launch vehicle that is designed in this evolutionary way could have a noticeably lower amortized development cost. There would still be some new development effort associated with the older parts to integrate them into the new vehicle and so the exact value of the developed structure-payload mass ratio might not be clear, but it would obviously be lower than the full value of the vehicle’s structure-payload mass ratio.

Even if the entire vehicle and its components are the result of a new development effort, the developed structurepayload mass ratio might still be lower than the vehicle’s structure-payload mass ratio if it uses identical stages. The simplest case of this is a bimese rocket where the booster and orbiter stages are identical. In the case of a bimese the developed structure-payload mass ratio would be half of the structure-payload mass ratio since only one of the stages would need to be developed and the production quantity for the stage doubled to provide the other half of the assembled vehicle. Unfortunately a bimese rocket suffers from staging inefficiencies that drive the total vehicle mass up above a conventionally sized rocket and cancels out the advantage of having the developed structure payload mass ratio be based on only half of the vehicle’s total mass. The hypothetical launch vehicle with a structure-payload mass ratio of 2 used to calculate current launch costs in Section III could be achieved as a two stage rocket and an average Isp of 380s if the lower stage is allowed to be 4.6 times larger than the upper stage. If both stages are forced to be the same size in order to achieve a bimese configuration, the structure-payload mass ratio increases to 4 and the developed structure-payload mass ratio still remains 2!

Trimese launch vehicles, composed of three identical major components with two components serving as boosters and one as an orbiter, were studied in the late 1960’s and early 1970’s and would have a developed structure-payload mass ratio that was only one third of the structure-payload mass ratio. The staging inefficiencies are not as bad with this design and using the same assumptions described above a trimese launch vehicle would have a structure-payload mass ratio of about 2.3 and a developed structure-payload mass ratio of about 0.8. The greater complexity of having one stage that could function as both a booster and orbiter would drive up the vehicle development structure cost, but this “back of the napkin” analysis suggests that a trimese configuration might still be a viable candidate for a low-cost next generation launch vehicle based on its low developed structure-payload mass ratio.

If amortized development cost can be reduced by using a trimese design, then it is logical to ask if using even larger numbers of identical stages might result in even greater savings. Ed Keith proposed12 an Asparagus-Stalk Booster composed of seven identical units that crossfeed propellant and are dropped in pairs as they are emptied. Using pressure fed rockets with an Isp of only 290s, this concept was project to have a structure-payload mass ratio of about 3.8 but a developed structure-payload mass ratio of 0.54. With the development of modular micro-electro mechanical rockets13, perhaps a launch vehicle could be constructed of thousands of identical subcomponents that detach when no longer needed. As the number of identical clustered stages in the vehicle increases, however, additional development problems occur that increase the developed structure cost and may wind up actually increasing the amortized development cost while the developed structure-payload mass ratio declines. This increased developed structure cost comes from the increasing technical challenge of accounting for the interactions between a large number of identical components and designing one stage that can function in any location in the vehicle despite the differences in the structural, aerodynamic, thermodynamic, and acoustic environments. These interaction problems would be reduced or eliminated, however, if the identical components that comprised the launch system could be physically separated from each other instead of clustered together. This is not impossible if a separated ascent stage space launch system is used, as in FLOC or Black Horse.

A separated ascent stage launch system is one where multiple vehicles launch and ascend separately, but cooperate through either aerial refueling or momentum transfer to help one of the vehicles attain orbit. Perhaps the earliest proposed example of a separated ascent stage launch vehicle is a variation of the Black Horse rocketplane. In addition to the normal mode of operation, a “speculative operational mode” was also proposed where the payload of the rocketplane could by increased by launching two rocketplanes simultaneously on a suborbital, exoatmospheric trajectory. Once out of the atmosphere, the rocketplanes would rendezvous and one would extend a refueling boom to the other and transfer its remaining propellant. The empty rockeplane would return to earth while the refueled rocketplane reignited its engines and continued to orbit.

Another example of a separated ascent stage launch system is the family of FLOC concepts developed by Alan Goff at Novatia Labs. The original FLOC space launch concept involved the simultaneous launching of a number of identical, reusable launch vehicles configured in a bimese arrangement. The bimese stages would crossfeed propellant and the booster stage would return to Earth when empty. The remaining single upper stages would pair up, dock with each other, and continue as before until the next set of boosters was empty. The remaining stages would continue this process of pairing up and crossfeeding propellant until only one upper stage was left with enough velocity and fuel to reach orbit.

The original baseline FLOC design used to illustrate the concept was based on a reusable 400 ton GLOW rocketplane with an Isp of 372 and a structure mass fraction of 29%. This structure mass fraction was a WAG† made by Novatia based on the Boeing 747’s dry mass fraction. With a maximum payload in the 32 unit flock configuration (31 boosters and 1 orbiter) of 178 tons this system would have an overall structure-payload mass ratio of 20.9 but a developed structure-payload mass ratio of only 0.65. This developed structure-payload mass ratio is only slightly better than the trimese configuration described above, but it is accomplished with 7s less of Isp and a structure mass fraction approximately three times as large. If the technical hurdles of mid-flight refueling can be overcome, then the FLOC vehicle would present a much easier design challenge than a reusable trimese vehicle.

The ability to increase the payload capacity of a FLOC launch system by doubling the flock size and adding another binary staging event to the trajectory gives it even greater insensitivity to weight growth than a conventional multistage rocket. If unexpected weight growth during vehicle development reduced the payload of a conventional launch vehicle below mission requirements then an expensive redesign effort would be required. With a FLOC launch system, the payload could be carried by the next highest flock size. Increasing the number of rockeplanes in the flock would increase hardware and operations cost, and require a redesign of the trajectory but it would not require the development of a new vehicle. It would not be nearly as easy to add a new lower stage to a conventional launch vehicle. The ability of the FLOC concept to change the number of binary stagings and the size of the flock to accommodate a wide variety of payload masses potentially allows the amortized development cost to be reduced even further by increasing the number of launches conducted. A twenty ton payload could be launched with an eight unit flock one day, and the next mission might use a thirty-two unit flock to launch a two-hundred ton payload. Launched singly or in pairs, the FLOC rocketplanes could also perform suborbital missions such as space tourism rides or micro-gravity experiments. By allowing one vehicle design to economically perform a wide variety of space launch missions the flight rate would be increased, with a corresponding improvement in its amortization factor. This capability also provides the opportunity to start with easier missions and build the FLOC fleet into a progressively more capable system as additional rocketplanes are acquired and the binary staging maneuver is perfected.

Rocket Economic Analysis

The FLOC vehicle analyzed in this section is a 400 ton rocketplane composed of 116 tons of structure, 71 tons of propellant, and 213 tons available for either payload or additional propellant. The configuration analyzed in this section will be the 32 unit flock launching its maximum payload of 178 tons to low Earth orbit. This configuration has a structure-payload mass ratio of 20.9 and a developed structure-payload mass ratio of 0.65. Despite the potential for the FLOC to increase the amortization factor with a wide range of payload capability, the amortization factor, a, used for this analysis will remain. Assuming the high structure mass fraction of the vehicle allows an inexpensive development program, as Novatia hopes, but that the vehicle development program cost is still within the historic range for rocket launch vehicles a developed structure cost of $20,000 per pound was initially chosen for the cost analysis.

This gives an amortized development cost per pound of payload delivered to orbit of (0.65 lb./lb.* $20,000/lb)/27, or $481 per pound of payload.

This is a considerable reduction from the $1,500 to $9,000 per pound specific amortized development cost estimated for current launch vehicles in Section II.

Assuming that the vehicle hardware costs are also at the low end of historic values for rocket launch vehicles at $1000 per pound of structure and that the reusable rocketplanes are worn out at a rate of 1% per flight, vehicle hardware costs were estimated to be $209 per pound of payload. The specific cost of risk for FLOC is more difficult to estimate than for conventional launch vehicles because a launch vehicle failure would not destroy the payload if it occurs on a vehicle that is not carrying the payload and because additional pairs of vehicles might be launched so that even if one had to drop out of the flock because of a malfunction the launch might still be carried out using a spare vehicle. Assuming a cargo value of $10,000 per pound and a chance of a vehicle failure during the launch of

1%, the cost of risk was estimated to be about $20 per pound of payload. This low probability of failure assumes that bringing two partially-fueled rocket vehicles together and docking them during ascent can be done routinely. If the binary staging maneuvers cannot be made safe enough to be carried out routinely then the FLOC concept will fail for reasons other than just economics. Propellant costs were estimated at about $4 per pound of payload, which is so small as to be within the error of other cost estimates in this section. Using a labor cost of $100 per hour, the specific labor cost would be $(2090*L) per pound of payload where L is labor intensity in units of hours per pound of structure. The lowest current launch vehicle labor intensities are around 1 hr/lb, which would give a huge specific labor cost for FLOC of $2090 per pound of payload because of the overall high structure-payload mass ratio of the 32 unit flock. One of the primary goals Novatia had for FLOC was to improve vehicle operations cost at the expense of a higher structure mass fraction. Assuming that in using the relatively high (for a rocket) structure mass fraction of 29%, Novatia succeeded in reducing the required labor intensity by an order of magnitude the resulting specific operations cost would be $209 per pound of payload.

The result of this rocket-based cost estimate, summarized in Table 4, shows that a cost to LEO of less than $1000 per pound of payload is probably achievable with FLOC if you neglect the cost of developing and demonstrating in-flight rocket refueling capability. This analysis also shows that the goal of radically reducing operations cost with a more robustly designed vehicle must be at least partially achieved to reach this. If operations labor intensity cannot be reduced below current best practices for space launch vehicles then the high structurepayload mass ratio of the concept will cause unacceptably high operation costs to overshadow any amortized development cost savings the system achieves.

Plane Economic Analysis

There is a shortage of historical data for estimating the developed structure cost of rocketplanes. The X-15 program cost $676 per pound of vehicle structure in then-year dollars, which converts to approximately $4,300/lb. in current year dollars. Based on data from http://www.astronautix.com, Scaled Composite’s SpaceShip One cost approximately $9,500 per pound of the rocketplane’s structure mass. The X-15 data is without the engine development cost, however the baseline FLOC design was sized for use with an existing engine design. In the future, information from the X-37 program may provide additional useful cost data. Using the higher example of $9,500/lb. and the amortization factor of 27, amortized development cost was estimated at $229 per pound of payload delivered to LEO.

Using a typical aircraft hardware cost estimate of $500 per pound of structure, the specific cost of hardware is estimated at $104/lb. Using high-end estimates of supersonic aircraft labor intensity of 0.01 hrs/lb. provides an operations cost of only $21/lb. Propellant and risk cost estimates were left the same in both rocket and plane cost estimate methods. The result of this airplane-based cost estimate, summarized in Table 5, shows a cost of less than $400 per pound of payload is probably achievable if you neglect the cost of developing and demonstrating in-flight rocket refueling capability. The key unknown factors that might alter this cost estimate are the amortization factor, which might be much higher than considering the versatility of the design, and the fraction of the vehicle expended per flight, which might be much lower considering the airplane-like structural mass fraction of the vehicle. If both of these unknowns were more favorable than the conservative estimates used in this analysis then specific costs around $100/lb. to LEO might be achievable with the baseline FLOC launch system in a 32 unit configuration.

The disparity between these two cost estimates brings up the question of whether a FLOC rocketplane would be more like a rocket or more like a plane in cost. FLOC would need to be capable of space operations and reentry, which are not airplane-like features. On the other hand in terms of structural mass fraction the FLOC baseline vehicle is closer to an airplane than a rocket, and might allow airplane-like design and manufacturing techniques to be used as Novatia hopes.

The proposed Black Horse rocketplane-to-rocketplane propellant transfer was to be done with a refueling boom extended between the rocketplanes during about a six-minute exo-atmospheric portion of their suborbital flight. With the propellant transfer expected to take about two minutes, that left approximately four minutes to conduct an orbital rendezvous between two cooperative spacecraft traveling on nearly parallel trajectories in the same direction. This is a much easier task than the rendezvous with a non-cooperative target headed in the opposite direction that anti-ballistic missile missions require.

---------------------------

This is indeed the crux of the matter

How much propellant you can transfer and how fast is going to be the key factor that drives everything else.
 

galgot

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Thanks for these explanations.

...

This would be a very high performance aircraft that can (barely) reach orbit and return than something that works well as a spacecraft. The Shuttle was a spacecraft, (and a pretty long duration one in orbital power and life support requirments) that could function as an adequate aircraft once back in the atmosphere. It's complex yes but in technical terms it's actually pretty straight forward as a design since what you need for, say, a hypersonic aircraft would work well for a suborbital/orbital spaceplane in general terms. It helps somewhat in that you don't actually have to design to the standards of a hypersonic cruise airframe but are only 'passing through' going up and down with your actual design realm being sub and supersonic flight.

It's not going to be cheap by any means but starting out with things like assuming a titanium airframe goes a long way towards addressing some of the problems.

Ok , but it eventually goes in orbit after being refuelled , anyway that is the goal, if i understood correctly.

...

Not sure what you mean by "autonomous atmospheric propulsion"? But OTHER than the refueling system you need to do about the same work for any other reusable concept so it's not qute that major of a change. Where things will get expensive is if you try and design and build to an 'aircraft' like standard as we've already seen. :)

By "autonomous atmospheric propulsion", i meant the jet engines used for takoff, then maybe used to fly back to a desired landing location once back in the atmosphere.
The Shuttle didn't had these, it was gliding once back in the atmosphere.
 
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Archibald

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Ok so I want to thanks AAm641 because he hit the nail on the head vis a vis the spreadsheet.

clear and imediate transparency, assuming full blame, read what follow

I screwed up. I knew Excel could be a bastard, and I wasn't wrong. More generally, Excel is a bastard if you're not gifted enough - one has to be very paranoid using it NOT to wreck the formulas.
My mistake:
there was a formula that linked many parameters together, and I killed it. This also explain WHY I wanted to share it with you, dear fellow forum members: I was quite sure one of you would pick any hole in it.

And AAM641 did it, so kudos and bravo to him. His reworked spreadsheet is fine.

So the imediate consequences are

- ok, 2-FLOC with keroxide and kerolox can't make it to orbit (freakkin' hell, was too good to be true)
The absolute best kerolox can do is the attached screenshot - with the usual caveats
- 3-FLOC with keroxide or kerolox can (but 3-FLOC less practical)
- 2- FLOC with hydrolox should be "safe" thanks to the vastly superior specific impulse but of course payload will take a hit.

The only good news: the mass of oxidizer to be transfered is now extremely small and tight. No more 80 000 pounds, that's the exact place were I bursted the Excel formula like an idiot. Barely 20 000 pounds.

As we say in french "100 fois sur le métier, remettez votre ouvrage" sigh
 

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steelpillow

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Some words about costs of this thing, taken from varied sources
Thank you for these snippets, very interesting. While I do not question the considerations they cover, the things they have left out lead me to doubt their conclusions. As one of them said,
FLOC would need to be capable of space operations and reentry, which are not airplane-like features.
Besides the refuelling system, also the hi-gee acceleration, in-space manoeuvring and re-entry protection would all add weight, cost and logistics penalties to say a 787 Dreamliner-like technology base. These all seem to have been passed over in those cost analyses.
Given that a 32-FLOC is only marginally cheaper to develop than a triamese with the same performance, one must question the practicality of getting 32 planes into the air within perhaps 5 minutes of each other and within a circle 100 miles across. The big plane formations of WWII required some planes to get off early and wait for the next lot to take off behind them: then, local timing issues might mean a loiter of 20 mins or more before the formation was sufficiently in place for it to set off and meet up with the outliers as it went. So there are both the operational risk of dropouts, requiring a few spare tankers to tag along, and the logistics of loitering requiring an early first-stage topup. And getting all that infrastructure sorted out comes within the initial development cost, never mind the subsequent operational costs.
Then there is the aircraft comparison issue in general. A spaceplane must have a very high wing loading if it is not to experience unacceptably high supersonic drag. This militates against economy at low speeds and against high-altitude airborne flight at subsonic speeds. Above say 30,000 ft (10 km) it will have to go supersonic, when refuelling will become impossible due to shockwave management (remember the SR-71/TAGBOARD experience) until say 300,000 ft (100 km) and enough speed, say 3,000 mph (5,000 km/h), is reached to buy it enough time. This places a strict limit on the number of refuelling cycles a FLOC could practically achieve. I'd suggest that say four would be a practical maximum, with 2^4 = 16, plus a couple of spares, so 18 as the maximum FLOC size - and even then, the first wave will just be a low-altitude subsonic top-up after everybody gets to the right place at last, so it is more comparable with an 8-plane FLOC in the referenced discussions. If any more were demanded, then the payload-carrier would not be able to make it from its second subsonic top-up to the first suborbital one.
I also recall the proposal to piggyback a mini-HOTOL off some carrier aircraft, back in the day. BAe found that a dedicated carrier aircraft would be too expensive to develop but that the 747 of the day was not quite big enough, so they opted for the Antonov An-225. Nowadays we might go for an Airbus A380 carrier with a SABRE powered orbiter, especially if we regard SABRE as an existing powerplant by the time we reach prototype build. This would reduce development costs to around those of a single-design biamese or flock approach. (OK I know some folks don't trust Alan Bond's arithmetic, but that is another question. BTW, he has just successfully tested his precooler at simulated Mach 5 airflow and temperature conditions).
So ultimately, I do not see the flock concept to be as potentially low-cost as those pieces suggest.

Somebody raised the issue of fuel flow momentum. A spacecraft does not have fine control of the throttle to counter it, especially if in ballistic free fall. Accelerating and decelerating a high-speed, high-volume flow would put significant tension on the boom. Flows within the craft could also cause powerful twisting moments around arbitrary axes at both ends. I think one might want to become adept at spraying red dye on buildings (a UK joke) before designing a practical system.
 

Archibald

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Yeah FLOC goes two bridges too far, really. As I said many times, beyond 4-FLOC is becomes quite difficult to synchronize the ballet. 8 is the very last end, above, it is just impossible.

The others snipets relate to Clapp rather than Goff, 2 to 4 (8 max) machines, and no docking, only brief refueling. I personally would stop at 4-FLOC cut into tanker+tanker and tanker+"future orbiter with payload". They first refuel in pair, then two tankers peel off the FLOC and return Earth, leaving a tanker+orbiter for a second refueling,before the orbiter goes into orbit (meh). Three refuelings overall. That's how it would work for 4-FLOC.
3-FLOC by contrast would be asymetrical: one tanker would refuel either a) the second tanker or b) the orbiter before returning Earth. The second tanker and orbiter would then perform a second refueling before the last vehicle press into orbit.

Well, as much as I red since 2007 about launch costs, my brain remain impermeable to it. Ranulfc as usual may do a better job than mine at explaining things.

Apart from that,

I have to recognize the Mk.1 keroxide bird really has taken a hit. Crap, if can't haul itself into orbit in 2-FLOC scheme but only in 3-FLOC, the gap with a single-bird carrying an expendable stage, is now a little too broad.
I mean, I liked it for its sheer simplicity (no cryogens, those kind of things) but 3-FLOC is a harder sell than 2-FLOC.

The only way to get back to a 2-FLOC scheme for mk.1 would be to cut into the PMF but I won't go this way because 0.85 is already pushing the limits for any rocketplane; and, the rocket equation being what it is (a very bad girl, to stay polite) every 1% of PMF toward 0.95 - SSTO would be paid in blood.
For your information, only methalox at 382 seconds (Elon Musk dreamed Raptor number) can make it to 9000 m/s with any meaningful payload. RD-0124 kerolox at 359 seconds fails.
 
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steelpillow

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I liked it for its sheer simplicity (no cryogens, those kind of things) but 3-FLOC is a harder sell than 2-FLOC.

How about modifying a solid-propellant + fluid-oxidiser system like the RocketMotorTwo >awful name< which powers the SpaceShipTwo? Make the rocket tube reloadable, like caseless ammunition? The "tanker" then just disgorges a nylon cartridge or two, which the orbiter picks up and eats. No stressed-out docking, boom or transfer valve required.
But there is still the nitrous oxide (oxidiser) to transfer. I note that its melting point is higher than the boiling point of liquid nitrogen, so at modest cryogenic temperatures (using N2 cooling) it would be solid and could be transferred as just another pellet. I wonder what its latent heat of fusion is? Too high and it won't melt easily enough, too low and it will melt before you can get it into the holding tank. Might need an insulating wrapper, or just a bit of extra cooling and get the journey done quick, like bringing ice cream home from the supermarket.
 

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Another alternative is to bit the bullet and proceed differently.
so 2-FLOC is dead for orbit, fine.

Sooo... 1-FLOC / single rocketplane with expendable upper stage still works. Well then, 2-FLOC with expendable stage should be able to boost a *larger* payload (but not the rocketplane itself !) into orbit.
Kind of intermediate step.
Bit the bullet of expendable stage for 2-FLOC and pitch the military "eventually, you can launch a rocketplane in orbit, with two tankers... well in many missions, F-16s refuel more than one time at more than one tanker.
Note that keroxide 2-FLOC certainly fails into orbit, but not by a large margin: very much like the Shuttle "Abort Once Around", that is, merely a single orbit before going down.
Still useful to perform a spying mission (of Russia ?) with a KA-80A camera... an incomplete orbit, sorta. Also useful for suborbital passenger transportation, P2P without entering orbit, perhaps with ricochet trajectories.

Mitchell Burnside Clapp once proposed a Star-48V for its Black Colt. Well, this might be a good idea to restore the 2-FLOC capability to boost a payload into orbit...

 
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Archibald

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I liked it for its sheer simplicity (no cryogens, those kind of things) but 3-FLOC is a harder sell than 2-FLOC.

How about modifying a solid-propellant + fluid-oxidiser system like the RocketMotorTwo >awful name< which powers the SpaceShipTwo? Make the rocket tube reloadable, like caseless ammunition? The "tanker" then just disgorges a nylon cartridge or two, which the orbiter picks up and eats. No stressed-out docking, boom or transfer valve required.
But there is still the nitrous oxide (oxidiser) to transfer. I note that its melting point is higher than the boiling point of liquid nitrogen, so at modest cryogenic temperatures (using N2 cooling) it would be solid and could be transferred as just another pellet. I wonder what its latent heat of fusion is? Too high and it won't melt easily enough, too low and it will melt before you can get it into the holding tank. Might need an insulating wrapper, or just a bit of extra cooling and get the journey done quick, like bringing ice cream home from the supermarket.

That's an intriguing idea, to say the least. Not sure the lower specific impulse of hybrid rockets allows it, though...

What puzzles me is, when compared to a very variety of fuels, how few choices there are for oxidizers bar the sacro-sanct LOX. And crazy-deadly mixtures like fluorine.
- H2O2
- N2O4 (the oxidizer side of storables, the other half being Mark Watney beloved hydrazine)
- N2O
...and that's it, goodbye and thanks for all the fish.
 

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Martin, let me start off with apologizing for taking this off on a false tangent. Re-reading the thread I realized I misinterpreted what you'd written and had taken your objections out of context and distored in my head, (and then on-line) what you'd said. My only excuse is that I was apparently much more ill than I thought and once my head cleared I now see where I took off on my side-trip. (Never post when you're running a high fever, even if you at the time don't KNOW you have a high fever...It was a fun weekend lets say :) )

Shall we move on with the discussion and try and forgive me for my mind slip :)

Randy
Randy,

no problem - I wasn't exactly my kindest, gentlest, most patient, best self. My apologies for that. Glad to hear you're doing better :).

Martin

Thanks, it's good to be back to only my normal case of meglomainia and delusions of Godhood ;D I think you were a lot more patient than I deserved, but thanks...

Randy
 

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Thanks for these explanations.

Ok , but it eventually goes in orbit after being refuelled , anyway that is the goal, if i understood correctly.

Yes it does but my point is unlike Shuttle/Buran it is more a high performance aircraft that can go into orbit than an actual spacecraft that can glide around in the atmopshere if it has to.

By "autonomous atmospheric propulsion", i meant the jet engines used for takoff, then maybe used to fly back to a desired landing location once back in the atmosphere.
The Shuttle didn't had these, it was gliding once back in the atmosphere.

Ok, and yes they are to allow a powered return to a compatable landing site, self ferry, and manuevering in the atmosphere. Keep in mind that both the orbital ground track of the destination and the reentry track can normally be far off from the intended ground base. As noted both Archibald and I got the impression that the original concept might have been aimed at servicing orbital tethers which are normally true-equatorial ground tracks so there's a LOT of over-water flying.

Randy
 

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Two (2) vehicles launch one with payload and one with propellant and fly to a sub-orbital, exo-atmospheric rendezvous, where they transfer propellant and one flies home while the other goes to orbit with the payload.

Yes, I am tempted to call this the basic scenario for suborbital refuelling or SOR. One might call it the SOR-LEO scenario.

I'd recalled the discussion on calling it SOR but in my bufflement couldn't find it again but... OK so we call it SOR for "Sub-Orbital Refueling (re-propellanting?) Which is a good start but we need something catchy, something that has meaning and will help 'sell' the idea....

Hmmm, the flight profile, trajectory, and mission are quite active, The propellant transfer is quite energetic, the manuevers needed spirited and lively. One could even say the whole concept was quite Dynamic, so we could call it ...

Oh, look, someone is pointing to the exit where my hat and coat lay in a burning pile... I wonder what that means?

Randy
 

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Sub-Orbital Aerial Refueling - SOAR, as in DynaSoar.

Problem solved.

Next issue ?
 

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when compared to a very variety of fuels, how few choices there are for oxidizers bar the sacro-sanct LOX. And crazy-deadly mixtures like fluorine.
- H2O2
- N2O4 (the oxidizer side of storables, the other half being Mark Watney beloved hydrazine)
- N2O
...and that's it, goodbye and thanks for all the fish.
It arises from the nature of carbon-chain fuels sitting on the reduction side of the chemical equation. The basis of the combustion engine is to put together a reducing agent which sucks oxygen and its kind out of things, with an oxidizing agent which releases oxygen or similar to attack other things, and then stand back while the heat released in the reaction blows everything out the exit door. It is really just the accident of having an oxygen-rich atmosphere which has led to the reducing agents being dubbed "fuel." Oxidants come from groups V, VI and VII of the periodic table, with the smaller atoms being the more chemically reactive ones. The group VII ones are the smallest in any row, on account of the extra pull on the electrons from their nuclei. So fluorine is the best oxidant, followed by oxygen and nitrogen. Then chlorine, sulphur and phosphorus. Large molecules with these elements in are rare and probably carbon-chain based anyway (though you can do something with sulphur), so we need small, simple ones that do not have a lot of reducing agents like hydrogen, lithium or carbon in them. Fluorine is too nasty to spew into the atmosphere, regardless of its potential. Thus, O2 is good, as are nitrogen oxides, but H2O or CO2 are not. H2O2 is a weirdo but works because it spits out an O at the slightest provocation, while N2H4 (hydrazine) is a reducing fuel, although both can be used as monopropellants under spontaneous decomposition. O3 (ozone) is also toxic and unstable, perhaps that is why it is not used. I think I have seen N2 proposed/used on occasion, but it is not as reactive/energetic as O2. The next row are almost as toxic as fluorine. It have probably forgotten something, but you should get some idea of why the list is so limited.
 
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RanulfC

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Ok so I want to thanks AAm641 because he hit the nail on the head vis a vis the spreadsheet.

clear and imediate transparency, assuming full blame, read what follow

I screwed up. I knew Excel could be a bastard, and I wasn't wrong. More generally, Excel is a bastard if you're not gifted enough - one has to be very paranoid using it NOT to wreck the formulas.
My mistake:
there was a formula that linked many parameters together, and I killed it. This also explain WHY I wanted to share it with you, dear fellow forum members: I was quite sure one of you would pick any hole in it.

And AAM641 did it, so kudos and bravo to him. His reworked spreadsheet is fine.

So the imediate consequences are

- ok, 2-FLOC with keroxide and kerolox can't make it to orbit (freakkin' hell, was too good to be true)
The absolute best kerolox can do is the attached screenshot - with the usual caveats
- 3-FLOC with keroxide or kerolox can (but 3-FLOC less practical)
- 2- FLOC with hydrolox should be "safe" thanks to the vastly superior specific impulse but of course payload will take a hit.

The only good news: the mass of oxidizer to be transfered is now extremely small and tight. No more 80 000 pounds, that's the exact place were I bursted the Excel formula like an idiot. Barely 20 000 pounds.

As we say in french "100 fois sur le métier, remettez votre ouvrage" sigh

Where did the concept lose it's viability and why? IIRC somewhere out there is the original Blackhorse spread sheet which showed a different result but I think that was due to size and payload of that vehicle?

How about modifying a solid-propellant + fluid-oxidiser system like the RocketMotorTwo >awful name< which powers the SpaceShipTwo? Make the rocket tube reloadable, like caseless ammunition? The "tanker" then just disgorges a nylon cartridge or two, which the orbiter picks up and eats. No stressed-out docking, boom or transfer valve required.
But there is still the nitrous oxide (oxidiser) to transfer. I note that its melting point is higher than the boiling point of liquid nitrogen, so at modest cryogenic temperatures (using N2 cooling) it would be solid and could be transferred as just another pellet. I wonder what its latent heat of fusion is? Too high and it won't melt easily enough, too low and it will melt before you can get it into the holding tank. Might need an insulating wrapper, or just a bit of extra cooling and get the journey done quick, like bringing ice cream home from the supermarket.

Hybrid rockets, (solid fuel-liquid oxydizer) tend to be less looked at and as always your propellant choices are important. (Primer: https://web.stanford.edu/~cantwell/...t_and_Rocket_Propulsion_Ch_11_BJ_Cantwell.pdf) while SS2 uses HTPB/N2O2 there was a bunch of work several years ago on wax/peroxide high thrust boosters, for use in ssto Mars Sample Return concepts. (https://aerospaceamerica.aiaa.org/year-in-review/hybrid-rockets-to-wax-or-to-whirl/)

Going to have to do a search for a second, and I found one (https://spacegrant.colorado.edu/COSGC_Projects/symposium_archive/2005/docs/115.pdf) Paraffin wax and H2O2 hybrid rocket motor. In general a higher regression rate of the solid fuel, (liquifying nad then aerosaling the propellant combination) gives better ISP and thrust and it turns out using a catalyst to decompose the H2O2 before injection into the fuel grain does wonders for all of it :) "Chill" the peroxide to around 41F/5C (Air Conditioning temps rather than cryogenic) will prevent if from decomposing indenfinatly so your main concern is the transfer and lock-down of the boosters for the next firing phase.
And I did not know this but apparently "gelled" peroxide is now a thing? (https://www.hindawi.com/journals/ijae/2018/5630587/)

A 30,000N/6,744lbf proposed hybrid design: https://www.researchgate.net/public...of_a_30kN_Paraffin-Based_Hybrid_Rocket_Engine

Randy
 

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If you have downloaded the previous spreadsheets, please crush them. I made a complete sweep of my PC and hard disk and nuked them into oblivion. There is no point in keeping flawed horse manure.

I may eventually removed them from the attachements.

The three below have been corrected. Basically - since propellant doesn't float free in the sky, you can't offload more prop than the tanker has remaining in its tanks. D'OOOH !!!

hqdefault.jpg
 

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Archibald

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I nuked the wrong spreadsheets in this thread. Sorry, but what is wrong... is wrong.
 

Archibald

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Basically 2-FLOC doesn't work if PMF is not 0.85 or more and specific impulse in the 350 range. No way any hybrid rocket can do it.

Now 4-FLOC 8-FLOC or crazy-FLOC, maybe...
 

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Basically 2-FLOC doesn't work if PMF is not 0.85 or more and specific impulse in the 350 range. No way any hybrid rocket can do it.

Do we have any figures for polyamide/N20? Wikipedia states that Branson has described it as "better" than HTPB but gives no specifics.
 

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Folks, I've found what to do with the failed 2-FLOC Mk.1 keroxide vehicle. The answer is Steve Pietrobon excellent N-STO paper.


In the spreadsheet I added calculations for a Star-48V solid-fuel kick stage, but Briz-M, Block-D, Agena, Centaur and a keroxide stage could be contenders.
 

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RanulfC

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Basically 2-FLOC doesn't work if PMF is not 0.85 or more and specific impulse in the 350 range. No way any hybrid rocket can do it.

Do we have any figures for polyamide/N20? Wikipedia states that Branson has described it as "better" than HTPB but gives no specifics.

Ya, I'm finding some polymer combustion studies but nothing that gives relative figures for performance. I'll keep looking.

Randy
 

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You will note that all the spreadsheets have the delta-v to Earth orbit fixed to 9000 m/s except for LH2 which gets 9300 m/s.

Now look at the attached document: I tried to reap the benefits of air-launch and dense propellants by criss-crossing Clapp, Sarigul Klinj, throwing into the lot Steven Pietrobon dense propellant delta-v calculations.
Basically hydrolox is 9300 m/s, the defunct keroxide is closer from 9000 m/s while kerolox is 9100 m/s.
Then, let's substract Clapp and Klinj air-launch calculations from all this

Capture.PNG

If that sounds convinving to you readers, then I may allow the Spreadsheets delta-v to Earth orbit to be cut from 9100 m/s to 8200 m/s. it makes some sense because it is one of Clapp original numbers for Black Horse.

What do you think ?
 

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I also recall the proposal to piggyback a mini-HOTOL off some carrier aircraft, back in the day. BAe found that a dedicated carrier aircraft would be too expensive to develop but that the 747 of the day was not quite big enough, so they opted for the Antonov An-225. Nowadays we might go for an Airbus A380 carrier with a SABRE powered orbiter, especially if we regard SABRE as an existing powerplant by the time we reach prototype build. This would reduce development costs to around those of a single-design biamese or flock approach. (OK I know some folks don't trust Alan Bond's arithmetic, but that is another question. BTW, he has just successfully tested his precooler at simulated Mach 5 airflow and temperature conditions).
Steelpillow, that's why I hope that Paul Allen's Scaled Composites Model 351 Stratolaunch (a.k.a. Roc) will survive after all :).

Martin
 

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that's why I hope that Paul Allen's Scaled Composites Model 351 Stratolaunch (a.k.a. Roc) will survive after all :).

H'mm, a two-stage spaceplane plus orbital refuelling of the second stage - a ROC FLOC perhaps? :D

Hey I was re reading Clapp black horse page where he explains that starting the rocket at Mach 5.5 cut delta-v to orbit to 20700 ft/s that is a mere 6300 m/s, saving 2700 m/s out of 9000.
This was written in 1994 when SABRE was merely a paper engine - REL was created in 1989 and it took until 2010 they made real progresses.

Earlier in the thread I mentionned MIPCC J58 and SERJ as possible improvements to a SOAR rocketplane. Well I would happily add SABRE to the list.
- civilian turbofan to mach 0.95
- military turbofan to mach 2.5
- MIPCC to mach 3
- J58 to mach 3.5
- SERJ to mach 4.5
- SABRE to mach 5.5
A gradual path to push airbreathing-to-rocket transition further and further; while keeping suborbital LOX transfer as a payload booster.

I was wondering if it would be possible to build some kind of hypersonic Mig-21 lookalike around a single SABRE. And indeed to drop that from Stratolaunch Roc. Or make that beast light enough Roc could carry two of them... and add a small suborbital LOX transfer between them to boost payload.

Incidentally one of the Skylon papers mention suborbital release as a possible payload helper.
A brief suborbital LOX transfer between two Skylons could get payload way above the present 37000 pounds.

As long as suborbital LOX transfer is mastered, the potential applications are far reaching and huge. Methane, kerosene or LH2 fuel will follow they don't need a transfer.

Same for the rocketplanes it doesn't matter if they are differents, what matters is the ability of performing suborbital LOX transfer to boost payload to Earth orbit.

Then imagine connecting that to a LUNOX industry coming from above... that's one of the reasons I want to fly the rocketplanes beyond LEO, to cislunar space, and even land them on the lunar surface IF Starship do it, achieving a lunar landing tail-sitter style.
 
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It might be possible to tweak one of the spreadsheets to include Skylon and see what gain could be achieved via a small suborbital LOX transfer between two of them.
Basically SABRE to Mach 5.5 cut the delta-v to orbit from 9000 to 6300, so put that later number in the case "delta-v to Earth orbit". Then enter Skylon C1 or D1 weight and prop loads, which are available from the many REL papers. The tricky part would be to assess how much LOX and LH2 is left at rocket transition, Mach 5.5; beyond that point Skylon is a dumb hydrolox booster, except reusable...
 

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I think the Roc's payload would depend on planned mission profiles and on the size/power of the production SABRE. Would a single-SABRE spaceplane without refuelling be able to carry out a useful mission? If not, then two engines would be necessary, with the options of putting them in one big spaceplane or two smaller ones plus refuelling. The single machine makes far more sense from the point of view of logistics and of technical risk. In terms of performance there must be a tradeoff between dragging a second engine into orbit or taking a longer time to accelerate on one engine, fighting gravity all the way. So I suspect that the larger plane would be most people's preference.
Once you have your optimal spaceplane for your basic mission, the sooner you can refuel then the bigger the payload you can get somewhere useful.
 

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It might be possible to tweak one of the spreadsheets to include Skylon

Bear in mind that Skylon reaches Mach 5+ while still in atmosphere. Say it reaches the Kaman line at Mach 5.5, it will be travelling ca 3,500 mph (for Mach 1 = 650 mph). By the time it gets high enough for a refuel meetup to avoid sonic shockwaves it will probably be travelling a thousand mph faster. The optimal point for a single refuel will be around half of orbital velocity, perhaps 8,000 mph. Theoretically three refuelling points would be the max needed to reach LEO if the tanks are fully refuelled at each point, at say 4,500, 9,000 and 13,500 mph.
 
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Archibald

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Ah I see, different trajectory due to airbreathing requirements. Must be pretty flat until Mach 6 and then zoom and boom upwards, on rocket power.

What I had in mind was a refueling at near orbital speed, derived from their own pitch for suborbital payload deployment.

"SKYLON has the capability of maximising the payload mass by performing a suborbital deployment and using a payload supplied propulsion system to perform the final orbit insertion manoeuvres. In this mission profile the vehicle flies an ascent trajectory which places it into an orbit that will allow a minimum of 5 minutes above an altitude of 135 km and achieve a flight path angle no steeper than minus 3 degrees at the re-entry interface of 120 km. The nominal transfer orbit to satisfy these constraints has been determined as follows:

Apogee: 157 km (radius of apogee = 6532 km)

Perigee: minus 2000 km (radius of perigee = 4375 km)

Velocity at apogee = 6966 m/s

Deployment is started 30 minutes after MECO, when the vehicle is above 135 km (a minimum deployment altitude determined by aerothermal constraints) the payload bay doors are opened and the payload is deployed. SKYLON then proceeds to separate from the payload, which is now independent, and closes the payload bay doors. After 2 – 3 minutes the payload fires its engine in order to raise it to its operational orbit. SKYLON then re-enters as with an orbital mission, and lands at a site about 10,000 km downrange from its launch site. The vehicle may then be “towed” back to its base.

The total payload mass that may be deployed during this mission is 30 tonnes which is determined by the structural strength of the payload interfaces. However, the payload must use some of this mass to provide its own propulsion system (or a dedicated propulsion stage) in order to raise itself to its operational orbit. Typical velocity increments which will be necessary are:

To 300 km circular orbit: delta-V1 = 858 m/s; delta-V2 = 42 m/s (2 burn Hohmann transfer) (author note 1: only 8000 m/s ?!)

I'm wondering if a second Skylon could fly along the first on such orbit, configured as tanker, the other one carrying, well, 30 mt instead of the usual 17 mt. The "gap" being filled by the tanker Skylon and a small LOX transfer. If that ever worked, then the payload gain would be well worth the pain.

Intriguing idea, really...
 

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Something to keep in mind about SABRE is it's actually NOT an 'airbreathing' engine design like a jet but a rocket-based propulion system that uses atmospheric oxygen through deep cooling the incoming air and feeding it into a rocket combustion chamber. It has turbomachinary to compress the air but that's about it for 'jet' parts. I bring this up because while Skylon planned to travel to around Mach-5 and something close to 200,000ft in "air-breathing" mode it would 'appear' (due to the deep cooling of the air and inlet) to the engine be traveling only around Mach-3 and about 100,000ft. It would then switch to internally carried LOX in the rocket and push on into space. MIPCC is somewhat similar in that it does much the same using water rather than hydrogen/helium and avoids the mechanical heat-exchanger by using direct injection. (The example would be the 'standard' F100 engine being capable of Mach 2 at 40,000ft without MIPCC but would be capable of Mach-4 at 70,000ft with due to the MIPCC keeping the compressor face at the Mach-2/40,000ft conditions)

So if launched from the ROC a "Skylon-ish" vehicle would fly to Mach-5 with a gradual rising trajectory and cutting in of on-board LOX till it reaches full 'rocket' mode where it continues at a slightly steeper trajectory, (to clear the atmosphere) until MECO. Note also that's going to be from LH2 depletion rather than LOX so you'd need to transfer LH2 since it uses 'less' (by design and therefore carries less) due to 'airbreathing' part of the way.

Part of the reason I brought up MIPCC is in an effort to avoid LH2 for obvsious reasons, but also for obvious reasons it won't work to feed a rocket engine the way SABRE will. Of course having said THAT I should note it will 'work' the same on some of the RBCC (Rocket Based Combined Cycle) systems like SERJ or an air-turbo-rocket system.

FLOC-ROC-SOAR?

Randy
 

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I'm not sure why people see an imbalance in Skylon's LOX vs LH2 refuelling needs. It will initially carry a small surplus of LH2 to last until it switches fully to onboard LOX, plus probably a further small reserve for low-altitude approach and landing. By the time it gets into space it is pure rocket with balanced fuel/oxidant consumption to match. All suborbital and subsequent refuellings will need to replenish those balanced LH2+LOX proportions, and only those balanced proportions, for pure rocket flight.
 

RanulfC

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I'm not sure why people see an imbalance in Skylon's LOX vs LH2 refuelling needs. It will initially carry a small surplus of LH2 to last until it switches fully to onboard LOX, plus probably a further small reserve for low-altitude approach and landing. By the time it gets into space it is pure rocket with balanced fuel/oxidant consumption to match. All suborbital and subsequent refuellings will need to replenish those balanced LH2+LOX proportions, and only those balanced proportions, for pure rocket flight.

Well one reason is we're not sure yet how LH2 propllant transfer is going to go, LOX we know we can do so it tends to be the baseline propellant to transfer. Second IIRC the Skylon was only using GH2 for the orbital manuever system and landing propulsion not LH2 so there wasn't really anything left after the initial burn? Of course we're not (I think) talking the actual Skylon since it would need a lot of modification to do a FLOC/SOAR type mission correct?

Oh and Archibald, not sure if you've seen this one?

Randy
 

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Nope, thanks for the link. ESA / europe RLV studies are much harder to find than the plethora of NASA NTRS ones.

I'm wondering the same about the O/F ratios.

My "stock" rocketplane weights 18 mt empty tanks, and 120 mt with the tanks full, leaving 102 mt, average.
Then at an O/F ratio of 7, kerosene represents 14 mt and H2O2, 84 mt, for a total of 98 mt.
Then switch to kerolox, and at an O/F ratio of 3: kerosene would be 25 mt and LOX, 75 mt.

I can't see why keroxide would need only H2O2 transfer while kerolox would take both fuel and oxidizer. I can't see 25 mt makes a difference big enough to justify fuel transfer. Just put a 25 mt kerosene tank, bit the bullet, and cut the LOX tank instead, filling the gap with a suborbital oxidizer transfer.
 
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steelpillow

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I'm not sure why people see an imbalance in Skylon's LOX vs LH2 refuelling needs. It will initially carry a small surplus of LH2 to last until it switches fully to onboard LOX, plus probably a further small reserve for low-altitude approach and landing. By the time it gets into space it is pure rocket with balanced fuel/oxidant consumption to match. All suborbital and subsequent refuellings will need to replenish those balanced LH2+LOX proportions, and only those balanced proportions, for pure rocket flight.

Well one reason is we're not sure yet how LH2 propllant transfer is going to go, LOX we know we can do so it tends to be the baseline propellant to transfer. Second IIRC the Skylon was only using GH2 for the orbital manuever system and landing propulsion not LH2 so there wasn't really anything left after the initial burn? Of course we're not (I think) talking the actual Skylon since it would need a lot of modification to do a FLOC/SOAR type mission correct?

Sorry, I must have missed it. How do we know that we can do LOX transfer in space?

AFAIK Skylon stores all H2 as LH2, there is no separate GH2 high-pressure tankage. Whether it is vaporised before reaching the thrust department is neither here nor there.

If the refuelling tank-and-boom system could be built as a self-contained cargo module able to mate with the standard fuelling ports, and dynamic flow forces were not problematic, then technically a standard suborbital Skylon could be refuelled without modification.
 

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Sorry, I must have missed it. How do we know that we can do LOX transfer in space?

Isn't the ISS replenished with LOX? I thought I'd read that somewhere but could be wrong that LOX was brought up on the Progress along with the RCS propellant.

AFAIK Skylon stores all H2 as LH2, there is no separate GH2 high-pressure tankage. Whether it is vaporised before reaching the thrust department is neither here nor there.

My information may be out of date :) IIRC it wasn't 'stored' per-se as much as what was left in the propellant tanks after the major manuevers were done. The tanks were big enough you had a lot of gas pressure to feed from.

If the refuelling tank-and-boom system could be built as a self-contained cargo module able to mate with the standard fuelling ports, and dynamic flow forces were not problematic, then technically a standard suborbital Skylon could be refuelled without modification.

True.

Randy
 
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