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Suborbital refuelling

Archibald

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Thanks for Ramon Chase documents. Very interesting. Particularly the 1991 paper. The usual consensus is that 100%-rocket-powered RLVs can't be winged and horizontal takeoff and landings, because rocket propellant is waaaaay too heavy. Notably the undercarriage. Ramon Chase however counter and destroy that argument, citing two notable examples. first is the B-58 undercarriage which was only 3% of its GLOW, with 1950 tech. Argument number 2 is Boeing RASV, although it needed a sled.

Note that Boeing RASV was the military offspring ( 1977-1982) of earlier NASA studies at Langley. They are available as Pdf at NTRS, and Chase actually briefly mentions them. It was called "Technology Requirements for Advanced Earth Orbital. Transportation Systems " and involved Boeing and Martin Marietta. it started in 1972 when Robert Salkeld and Len Cormier picked the curiosity of NASA Langley Gene Love which had his own pet concept, the "Continental/SemiGlobal Transport "
So in a nutshell
1971 Len Cormier Windjammer
1972 Gene Love CGST
1973 Robert Salkled "earth to orbit shuttle" studies
1975-1978 Boeing / Martin Marietta / Langley studies
1976-1983 Boeing RASV

Where it is extremely interesting is that Robert Salkeld, at some point, introduced rocket-powered HTOHL designs but they were so heavy, he proposed two solutions. Option 1 was the sled and led to the RASV. Option 2 was subsonic refueling, 20 years before Mitchell Burnside Clapp. Option 2 started with Salkeld circa 1973 and then found its way across the Langley studies.
And it was rejected.
Why ?
Because SSME - and once again, the disastrous influence of the very flawed Space Shuttle.
RASV was powered by a pair of SSMEs (and incidentally, back in 1971 Cormier Windjammer had SSME ancestors, ISINGLASS XLR-129). those engines are hugely powerful and swallow colossal amounts of deep cryogens - LOX and LH2. Plus the later abysmal density.
And on top, another pernicious influence of the Shuttle was a requirement for a 65 000 pounds payload - for the Langley studies. This was dropped by the military for the RASV.
All this ruined any hope for subsonic refueling: as rightly noted in the Langley studies, a colossal, expensive and dangerous tanker aircraft would be needed.
By contrast, note that when Clapp "reinvented" subsonic refueling 15 years later (1978 - 1993) the Black Horse was
- non deep cryogen, keroxide instead
- no SSME huge power
- extremely light and small, no bigger than a F-16
- screw any Shuttle legacy - no SSMEs, no 65 000 pounds payload.
This made the tanker shrinks from a monster cryogen carrier to a merely modified KC-135Q of SR-71 fame.
Also note that the RASV was related to Bernard Schriever SAMSO and maybe to Safeguard, because they mentionned Grand Forks AFB and "B-52 -like ground ops". But with SSME and deep cryogens ? no way.

I have a sneaking suspicion Clapp knew of RASV through his USAF background and also through the 80's TAV studies - across the 90's Black Horse is repeatedly called TAV.
There are probably many indirect but very real legacies between RASV and Black Horse. I think their aerodynamic shapes, for a start, were identical. Plus Boeing pushed RASV again in the early 90's for DC-X / DC-Y / X-33 (from memory).

Chase 1991 paper counter-intuitively shows that a rocket-powered HTOHL vehicle, far from being a five-legged-sheep, would actually have some intriguing advantages. Look at the scores at the end of the paper: at 25, it win the day ! And just like Clapp two years later, Chase uses a very real piece of metal as basis for its claim: that bit of RASV structural test article.
So thanks for the "Chase tip" Martin Bayer. While SOR might not be the definitive end of it, at least we agree on one thing: rocket powered HTOHL, if a trick can be found to get it off the ground (bimese or TSTO or sled, subsonic or suborbital refueling or maglev or something else) is well worth the effort.
 
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martinbayer

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Archibald,

while, based on rocket engine life limitations as well as inflight reignition risk considerations, my mind is rather firmly made up with respect to RLV landing modes, I continue to remain fairly agnostic on launch modes, be they (in alphabetical order) air launch, ground accelerator launch, horizontal runway launch, or vertical launch (and I'm fairly confident that I left out a good number of more arcane alternatives). The optimum launch mode will IMHO ultimately depend on individual mission parameters, system requirements, geographical conditions, available infrastructure, and technology readiness, and may very likely differ on a case by case basis. As far as main propulsion goes, you'll need rockets at some point during ascent anyway, so you better have a darn good justification, like for example achieving launch cross range and different orbit inclinations from a single launch site, to complicate things by throwing any kind of airbreathers in the mix. I am however just old fashioned and ornery enough to insist that my favorite RLV be sent up in one single unified compound rather than literally be composed on the fly, because that allows you to verify and, if necessary, fix all critical interfaces on the ground, thank you very much - as opposed to that old Neil Sedaka song title, breaking up is not hard to do at all, while there is a reason why aerial trapeze acts are considered circus thrills (and I'm looking at you as well, rotating orbital tether) ;).

Martin
 
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steelpillow

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circus thrills

The great airships often had a walkway the length of the upper hull. Crew could emerge into the extremely thick boundary layer at the tail and walk comfortably most of the way forward. Only on aeroplanes is going outside a circus act.

Back in the pioneer days designers learned to make aeroplanes aerodynamically stable. That became the norm, with unstable designs reserved for extreme performance such as interceptor, aerobatic and racing types: circus acts, if you will. Then fly-by-wire came along and nowadays even commercial airliners can have relaxed static stability.

SpaceX is busy proving that routine inflight reignition is not so hard, while Reaction Engines has demonstrated all the key breakthrough technologies for the prototype airbreathing rocket it is now seeking funds to build.

So one technology's circus thrill is another's bread and butter. Although I too am old and ugly enough to be sceptical that sub-orbital flock or orbital tether techniques can ever find a viable niche, they both appear technically possible and I like to keep an open mind.
 

Archibald

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nice. This thread started all wrong but ended on a better note.
Nota bene: I don't know why people somewhat assume I come as some arrogant prick utterly convinced that suborbital refueling is like bread coming in slices, a revolution making everything else obsolete, bla bla, you get the point.

This has NEVER been my intention, attitude or my tone.

I see some explanations to this. First english is not my native language. Secondly - internet trolls, we had plenty of them at Nasaspaceflight.com - pipe launcher, swala, many others...
 

steelpillow

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This has NEVER been my intention, attitude or my tone.

If I may say so, I would suggest that in some of your posts, especially the earlier ones, enthusiasm + language barrier = overstating your case.
Human nature being what it is, this encouraged your critics to overstate theirs.
If I was one of those over-energetic critics then I must apologise.
I too am very glad that things have now cleared up; hope that this would happen was my main reason for bringing the discussion to its own thread.
 
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martinbayer

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Hello Steelpillow,

in my best theoretical understanding, trying to venture outside of a launch vehicle during powered ascent would not just be a mere (if utterly amazing) circus act, given the acceleration impact, but even if you physically were able to do so, guaranteed suicide, considering the various other forces acting in the particular environment. My "circus act" reference pertained to the concept of two vehicles at high velocity and altitude having to rendezvous and mate with each other in a space and time critical window to transfer chemically highly reactive fluids from one to the other in short order while in free fall.

I fully acknowledge and appreciate the progress that SpaceX, Reaction Engines, and others have made in RLV technology, but being able to pull off a particular technological feat doesn't automatically mean it's the best way to go forward. In terms of rocket engine reignition, there is still a notable difference between "not so hard" and "almost risk free", and even when it works, you still fritter away precious engine burn life that you would in my view much better spend on ascent and acceleration cycles rather than on repeated deceleration, descent, and landing maneuvers in the Earth's atmosphere.

I try to keep an open mind as well, but over the decades I've seen lots of RLV trade studies with fairly consistent conclusions and results to inform my view on what's merely possible versus what's probably optimal. Nevertheless, I'm with you - show me that it works *and* that it is a superior solution over existing technologies, and I'll happily switch my mindset.

As an aside, am I correct in assuming that your online moniker is an indication that you are an acolyte of Karel Bossart?

Martin
 
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steelpillow

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Hello Steelpillow,

in my best theoretical understanding, trying to venture outside of a launch vehicle during powered ascent would not just be a mere (utterly amazing) circus act, given the acceleration impact, but even if you physically were able to do so, guaranteed suicide, considering the various other forces acting in the particular environment. My "circus act" reference pertained to the concept of two vehicles at high velocity and altitude having to rendezvous and mate with each other in a space and time critical window to transfer chemically highly reactive fluids from one to the other in short order. I fully acknowledge and appreciate the progress that SpaceX, Reaction Engines, and others have made in RLV technology, but being able to pull off a particular technological feat doesn't automatically mean it's the best way to go forward. In terms of rocket engine reignition, there is still a notable difference between "not so hard" and "almost risk free", and even when it works, you still fritter away precious engine burn life that you would in my view much better spend on ascent and acceleration rather than on deceleration, descent, and landing in the Earth's atmosphere. I try to keep an open mind as well, but I've seen lots of RLV trade studies with fairly consistent conclusions and results over the decades to inform my view on what's merely possible versus what's probably optimal. Nevertheless, I'm with you - show me that it works *and* that it is a superior solution over existing technologies, and I'll happily switch my mindset.

As an aside, am I correct in assuming that your online moniker is an indication that you are an acolyte of Karel Bossart?

Martin

Hi Martin,

Yes, I am well aware of your context for the "circus" remark. My point was a very general one, that a "circus act" for one class of vehicle may be humdrum routine for another. On the other point, I am not convinced that "almost risk-free" can yet be applied to any space technology. The realisation that, actually, a little extra fuel and tankage for landing can be more optimal than all the traditional workarounds to avoid the need for them, is one of the keys to the low cost of a SpaceX launch: if their costs were on the high side I would be a good deal less impressed. But if you want sufficient track record to demonstrate superiority in the field, we shall have to wait for Elon Musk and his ongoing programme. He will obviously have to make sure that he does not run out of fuel again when he has passengers on board, but my main concern for its future is that it tends to burn harmful kerosene not clean hydrogen, but that is another story.

My alias comes from my love of surrealism, blended with an unwanted web domain left over from a company who literally made steel cushions by inflating welded sheets with compressed air. There is also a habit in parts of Africa, of carving made-to-measure beds from solid wood, with the wooden pillow set on a stalk. They are said to be pleasantly cool and very comfortable, as long as you lie still. Had I had aerospace in mind, I might have envisaged a surreal cyberpunk metal-clad airship. In fact, now you have got me thinking - sheet steel would be tougher than the usual aluminium and a little positive gas presure a la Bossart would stiffen it nicely... ;)
 
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Archibald

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To put matters in perspective:
average numbers are
- 50 000 pounds of LOX or H202 (no fuel)
- 4 minutes for the transfer
- closest point of reference : KC-10 and KC-46 tankers, 1500 gallons of kerosene per minute. Roughly 4500 kg or 10 000 pounds.

Dare I ask - what could go wrong ?

Instabilities during transfer ? CG shifts ? something else ?
 

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"Slightly" incorrect
I think you will find my account perfectly correct.

Intent matters and in the case of the Shuttle as a TAOS concept:
1) It's designed to carry and bring back the crew and possibly a payload from Earth orbit
2) The ET is staged almost at orbital speed
3) Since making the ET reusable would entail a major loss of payload margin it is designed and intended to be expendable
4) Since you want to reuse the engines you therefor have to mount them on the Orbiter and return them with the crew/payload

It's the "SpaceX-Upper-Stage-Recovery" problem in nature in that you go from a light weight expendable external tank to a heavy and robust booster stage which wasn't compatable with the TAOS, (even though NASA kept hoping it would be) concept.

The Energia had the same issue in that while they could consider recovery of the booster stages the main stage and engines would end up needing to be built both more robust and with an added reentry and landing system which was found to impact the overall payload enough it was never seriously considered as an option.

Randy
 

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RanulfC

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"Slightly" incorrect
I think you will find my account perfectly correct.

Intent matters and in the case of the Shuttle as a TAOS concept:

Well, I am not sure whether you are commenting on my intent or NASA's, but either way I am not aware that I am disagreeing with anything except your intent to disagree with me. Can we please move on before this gets any more complicated? :rolleyes:

NASA's actually but no problem :)

Randy
 

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I am attaching a spreadsheet with my calculations.

Thanks to you, again. This one was a Godsend.

So, latest developments in my pet peeve crusade are pretty exciting. Because I'm locked home, I have a lot of time in my hand to explore that rabbit hole.

So on one side is Clapp, suborbital refueling (SUOR)

On the other, is Alan Goff suborbital docking (SUOD)

Both with 8 vehicles. Why eight ? because that's the only place where these two overlap - SUOR maxes there, SUOD... soars.

Crucially,
- SUOD doesn't work for less than 4 vehicles, and even there it needs hydrolox and 455 seconds specific impulse. Its most interesging PRO: docking is a little less dangerous, and faster, than prop transfer.
- SUOR works right from 2 - 4 vehicles, with non-hydrolox props. Mirroring SUOD, it is more dangerous but far more efficient.

Basically, with SUOR you transfer raw fluid. With SUOD you transfer weight of the second vehicle, now attached to the first.
(and yes, I avoided *SOD* which is an extremely offensive and very unfortunate accronym ROTFL)

End result: with 8 vehicles, SUOD can haul 62 mt and SUOR, 86 mt.
 

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  • Suborbital_REFUELING_8_vehicles.xls
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Archibald

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Critic of bimese and trimese vehicles. From the attached paper

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 structurepayload
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.
 

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martinbayer

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Critic of bimese and trimese vehicles. From the attached paper

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 structurepayload
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.
Taylor does not provide any reviewable quantitative engineering comparison at all, not even in the form of that “back of the napkin” analysis, to substantiate these assertions. I call BS.
 

steelpillow

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Financial cost equations are all very well, but without worked examples they are meaningless. One has to remember that the Shuttle was kinda quadramese for launch and biamese for orbital acceleration, and some SpaceX configurations not only have far more commonality between components but also drastically cut costs. I see no recognition of such basic realities in that analysis.
 

Archibald

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some counter-intuitive stuff about air-launch.

According to Marti Sarigul-Klijn
- optimal air launch, independant of the speed of the mothership, is 45 000 ft and 30 degree angle-of-attack.
Once there
- Mach 1 mothership substracts 1100 m/s out of 9200 m/s to Earth orbit. Fine.

One would think that, since Mach 2 is 200% of Mach 1 (*2) and Mach 3, 300% (*1.5), the "air launch boost" should rise accordingly.

Well, no... it rises, but slower.

1100 m/s * 2 should substracts 2200 m/s on the way to Earth orbit. Instead, it is 1600 m/s. That is, not *2 but *1.45

1100 m/s *3 should substracts 3300 m/s or, alternatively: Mach 2 to Mach 3 is *1.5 and 1600 m/s *1.5 should substracts 2400 m/s.
Well, no. The "air launch boost" is "only" 2000 m/s. That is, *1.25

So much for the pain of launching from a Mach 3 aircraft, rather than Mach 2 or even Mach 1...


So the message is: when getting a faster mothership,

The "air launch boost" grows SLOWER that the mothership speed gain (Mach 1 to Mach 2 to Mach 3).

For what reason, I don't know.

Probably because flying hypersonic is a bigger PITA than flying supersonic, itself a bigger PITA than flying subsonically. Just ask Concorde and SST why did they failed against the 747... the deeper into the supersonic and hypersonic regimes, the harder. That's the crux of the matter, and also aplies to air launch - even before "separation issues" are considered (hint: D-21 hitting a SR-71 mothership at Mach 3 with the two having a very bad day).
 

steelpillow

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Mach 1 at 45,000 ft is only 300 m/s so I am very suspicious how it can give a 1100 m/s saving. If we assume a prograde equatorial flightpath, that adds only 460 m/s, giving a total boost of 760 m/s. There is another hidden assumption in those numbers, or a mistake. Do you have a link to the source calculations?
 

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And I can't find the Purdue link to that goddam paper. It wasn't the classic Pdf and now I can't find it anymore.
 
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RanulfC

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some counter-intuitive stuff about air-launch.

According to Marti Sarigul-Klijn
- optimal air launch, independant of the speed of the mothership, is 45 000 ft and 30 degree angle-of-attack.
Once there
- Mach 1 mothership substracts 1100 m/s out of 9200 m/s to Earth orbit. Fine.

One would think that, since Mach 2 is 200% of Mach 1 (*2) and Mach 3, 300% (*1.5), the "air launch boost" should rise accordingly.

Well, no... it rises, but slower.

1100 m/s * 2 should substracts 2200 m/s on the way to Earth orbit. Instead, it is 1600 m/s. That is, not *2 but *1.45

1100 m/s *3 should substracts 3300 m/s or, alternatively: Mach 2 to Mach 3 is *1.5 and 1600 m/s *1.5 should substracts 2400 m/s.
Well, no. The "air launch boost" is "only" 2000 m/s. That is, *1.25

So much for the pain of launching from a Mach 3 aircraft, rather than Mach 2 or even Mach 1...


So the message is: when getting a faster mothership,

The "air launch boost" grows SLOWER that the mothership speed gain (Mach 1 to Mach 2 to Mach 3).

For what reason, I don't know.

Probably because flying hypersonic is a bigger PITA than flying supersonic, itself a bigger PITA than flying subsonically. Just ask Concorde and SST why did they failed against the 747... the deeper into the supersonic and hypersonic regimes, the harder. That's the crux of the matter, and also aplies to air launch - even before "separation issues" are considered (hint: D-21 hitting a SR-71 mothership at Mach 3 with the two having a very bad day).

Which paper was it in? If I can find it in my numerous binders I'll see if I can scan it in or find an on-line copy.
(I suspect it's this one, https://arc.aiaa.org/doi/10.2514/1.8634, or it's 2004 cousin. First is "Air Launch Earth to Orbit: Effects of Launch Conditions and Vehicle Aerodynamics" (on-line copy here: https://www.researchgate.net/public...of_Launch_Conditions_and_Vehicle_Aerodynamics) and the second, (actually first :) ) "Air Launching Earth-To-Orbit Vehicles: Delta V gains from Launch Conditions and Vehicle Aerodynamics" (AIAA 2004-872)

Mach 1 at 45,000 ft is only 300 m/s so I am very suspicious how it can give a 1100 m/s saving. If we assume a prograde equatorial flightpath, that adds only 460 m/s, giving a total boost of 760 m/s. There is another hidden assumption in those numbers, or a mistake. Do you have a link to the source calculations?

Note that Figure 1a in the first paper on-line version (launch velocity of 340m/s or around Mach 1 and an angle of between 30 and 60 degrees) shows a Delta-V gain of between around 900 to almost 1200 m/s. That may be where the numbers came from. And please note that all launch data is assumed at varying speeds but at sea-level for launch purposes.

Randy
 

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Mach 1 at 45,000 ft is only 300 m/s so I am very suspicious how it can give a 1100 m/s saving. If we assume a prograde equatorial flightpath, that adds only 460 m/s, giving a total boost of 760 m/s. There is another hidden assumption in those numbers, or a mistake. Do you have a link to the source calculations?
Note that Figure 1a in the first paper on-line version (launch velocity of 340m/s or around Mach 1 and an angle of between 30 and 60 degrees) shows a Delta-V gain of between around 900 to almost 1200 m/s. That may be where the numbers came from. And please note that all launch data is assumed at varying speeds but at sea-level for launch purposes.
I don't care where the figures came from, they do not stack up.

Next, note what Archibald actually posted:
According to Marti Sarigul-Klijn
- optimal air launch, independant of the speed of the mothership, is 45 000 ft and 30 degree angle-of-attack.
Once there
- Mach 1 mothership substracts 1100 m/s out of 9200 m/s to Earth orbit. Fine.
So "please note", as you so quaintly put it, that 45 000 ft ... once there is not exactly sea-level.
 

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I finally found the silly thing (amid a load of horse manure by that idiot RGclark)



Marti Sarigul-Klijn is hardly a nobody. I also have a different paper by Kirk Sorensen that also note similar benefits (now I have to find this one, ROGNTUDJU)

Did I mentionned that RGclark is the most stupid, annoying troll, ever ?
 

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I finally found the silly thing (amid a load of horse manure by that idiot RGclark)



Marti Sarigul-Klijn is hardly a nobody. I also have a different paper by Kirk Sorensen that also note similar benefits (now I have to find this one, ROGNTUDJU)

Suspect this one?

"Crossbow Air-Launch Trade Study"

Did I mentioned that RGclark is the most stupid, annoying troll, ever ?

No, really? I got the impression he's one of your favorite internet "experts" :D

I don't care where the figures came from, they do not stack up.

So the figures are wrong?

Next, note what Archibald actually posted:
> <
So "please note", as you so quaintly put it, that 45 000 ft ... once there is not exactly sea-level.

That's why I pointed that out but it also notes that sea-level is a lousy place to 'air-launch' a rocket but it was a basis for the calculations to match the already existing Minotaur flight profile. The report posted itself states that sea-level drag is a major factor in the calculations and that for example:
"The magnitude of the benefit depends on the LV’s flight-path angle at engine start. For example, examine Fig. 1a and consider a 340-m/s (Mach 1) launch at a 60-deg flight-path angle at a sea-level altitude. Although this is an unrealistic launch condition, note that the Vgain is 560 m/s. The gain comes from both the initial launch velocity (340-m/s) and a change in the losses. In this case, the high velocity near the ground will increase drag losses, but the same conditions decrease gravity and steering losses even more."
(Page 571 paragraph "Velocity")

The cited/posted pdf that Archibald put up says:
"A study by Klijn et al. concluded that at an altitude of 15,250 m, a rocket launch with the carrier vehicle having a zero launch velocity at an angle of attack of 0° to the horizontal experienced a Δv benefit of approximately 600m/s while a launch at a velocity of 340m/s at the same altitude and angle of attack resulted in a Δv benefit of approximately 900m/s.

The zero launch velocity situations can be used to represent the launch from a balloon as it has no horizontal velocity. Furthermore, by increasing the angle of attack of the carrier vehicle to 30° and launching at 340m/s, they obtained a Δv gain of approximately 1,100m/s. Increasing the launch velocity to 681m/s and 1,021m/s produced a Δv gain of 1,600m/s and 2,000m/s respectively."
(Page 387)
(And note that 15,250m is actually over 50,000ft not 45,000ft)

The Klijn paper notes that at altitudes above 12,000m (over 39,000ft) the air density and it's effects are about one-quarter of that of sea-level with a similar decrease in drag and increase in engine performance. (Page 571 paragraph "Altitude") So with the assumptions that 30 degrees is in fact more efficient than a 60 degree AoA, (again the graphs point this out) is it so hard to see that your Delta-V gain could go from around 560-m/s to around 1100-m/s? (It actually shows better than that at 30 degrees and Mach 1 AT sea-level)

Randy
 

Archibald

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Folks,
I have learned about a near-perfect oxidizer for my scheme: NYTROX.

One interesting aspect of suborbital refueling relates to oxidizers. Provided the O/F ratio is above 6, it becomes possible to transfer only oxidizer and not fuel.
-H2O2/RP-1 O/F ratio is 7 - good.
-N2O/ RP-1 O/F ratio is 7 to 9 (!!) even better.
-LOX/H2 O/F is 6, good enough.
...
-Unfortunately LOX/RP-1 ratio is 2.3 so kerosene would have to follow LOX; a two-fluids refueling system, including LOX cryogen... ugh.

Overall, when you think about it rocket oxidizers are a little discouraging. There is LOX, and there is... next to nothing.
Exactly:
-H2O2
-N2O4
-N2O
And that's it, goodbye and thanks for all the fish.

What's more, all four of them are PITA as far as suborbital refueling is concerned.
- LOX is -183°C cryogen
- N2O4 is nasty, toxic, corrosive stuff
- H2O2 and N2O are a little nicer but they have their own explosive issues (Virgin Galactic 2007, cough, HMS Syddon, cough cough)
And their performance is not great.

Now nytrox... I kind of like it. The gist of the idea is to blend LOX and N2O to try and get the best of both worlds. Well, as far as suborbital refueling is concerned, it's a Godsend.
- with RP-1 it drags the O/F ratio away from LOX (2.3) and closer from N2O (7+)
- it solves LOX temperature issues, making it "warmer" (-50°C average, depends from the mix)
- with RP-1 it solves N2O shit specific impulse of 310 and drags it toward LOX (330 - 360) - and thus better than N2O4 & H2O2 (320 at best)
- it solves N2O explosive nastiness that killed 3 Virgin workers in July 2007

All hail nytrox/RP-1 ! By this metric, the "ideal" engine for my rocketplane might be a RD-0124 (Soyuz 2 upper stage) running on syntyn and nytrox. A begnin oxidizer (not explosive, not deep cryogen, not storable-toxic) with an ordinary fuel (kerosene) with a good enough specific impulse in the 340-345 range.
 
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