On Henson's observation "..The approach gets gets around the ........the poor mass ratio of conventional rocket or rocket planes."
In 1964, work in the US suggested 30% payload mass fraction for airbreathing SSTO aerospaceplanes that collect their entire lox requirements in high speed flight while ascending to orbit. This contention was independently verified and published in India since 1989, with emergence of the "Hyperplane" concept. India's concept studies on SSP centre on this type of "aerobic" reusable launch vehicle.
High temperature light weight front-end heat exchangers for flight to Mach 7 is a critical technology. Japan reported development of such HEX tubes in 1987 or '88, tested to Mach 9. Materials for uncooled scramjet combustors is another. NASA reported studies in this area..
"Silo-mentalities" within and between nations are a severe constraint to progress. An integrated , cooperative, global approach to an SSP-RLV system-of-systems design and mission is inescapable.

<DIV>For,</DIV> <DIV>Dana Andrews</DIV> <DIV></DIV> <DIV>Thanks, Andrews, glad you recall our presence at the AIAA meet.I shall be happy to keep in touch with you off-line. My e-mail is gopalavatar@yahoo.co.in</DIV> <DIV></DIV> <DIV>Just a point in clarification :take for instance a 275-tonne aerobic vehicle that takes off with zero lox on board.It tanks in almost 200 tones of LOX while flying at 26-30 kms at speed increasing fro Mach 3.5 to Mach 7/8 in about half-an-hour. </DIV> <DIV></DIV> <DIV>The Hydrogen Fuel at take-off (60%includes4 % residuals in orbit ) = 0.6 x 275 = 165 tonnes</DIV> <DIV>Fuel consumed take-off to orbit = 0.56 x 275 = 154 tonnes</DIV> <DIV>Hence mass in Orbit = Dry structure + payload + Residuals = 275- 154=121 tonnes</DIV> <DIV>Payload = 25 tonnes</DIV> <DIV>Hence Dry Structure weight (with Fuel residual for return) = 121-25 = 96 tonnes</DIV> <DIV>Ratio, Payload/Dry Structure (+ residuals) = 25/96 = 26%</DIV> <DIV>Ratio: Payload to take-off weight = 9.1% for a take-off mass of 275 tonnes</DIV> <DIV></DIV> <DIV>Our study shows that with very high take-off weight, this payload fraction increases. Graphical extrapolation of our results show that the 30% figure quoted by John Hopkins may have been obtained when Take-of weight is 1000 tonnes. However the paper referred to does not give this figure</DIV> <DIV></DIV> <DIV>Our studies also show that this aerobic concept with liquefaction at high speed has another unique advantage , it enables scaling down without loss of orbital capability but with reduced payload fraction. At 25-tonne take-off weight, the size of a fighter aircraft, the take-off weight is 25-tonnes to deliver 1-tonne in orbit ie 4% of take-off weight</DIV> <DIV></DIV> <DIV>Given very great 'realism' in design premises, i.e even with heavier airframe structures, airbreathing engines and reduced specific impulse , we should be able to get 1-tonne in orbit with, say, a 50-tonne take-off weight. Still far , far less risky and less expensive to try out than the hundreds of tonnes jumbo-size vehicles proposed so far.</DIV> <DIV></DIV> <DIV>The whole trick lies in the heat exchanger and this is where you have excelled. And of course, the aerospacevehicle system design and performance measurement tools used for multi-variable optimization, like what we have custom-built and used.</DIV> <DIV></DIV> <DIV>Regards,</DIV> <DIV>Gopal

--- On Thu, 18/12/08, IntenseDebate Notifications <notifications@intensedebatemail.com> wrote:
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Hence I wanted to speak to you urgently at the AIAA event in 2001, that all published efforts in the US post Challenger ie from 1986 onwards, including yours, have studied air collection and oxygen separation only in the sub-sonic regime under the (misconstrued) premise that this was the region of greatest air density and dynamic pressure; and hence most conducive to lowest weight of the ACES technology.

Unfortunately, this error in thinking in choice of flight regime for ACES persisted for several years until 1995 Reference Vandenkerchove J and Czysz P “ SSTO Performance Assessment with In-flight LOX Collection”, Acta Astronautica Volume 37, 1995, Pages 167-178, brought out the need for US to redirect work on lox collection to hypersonic speed regimes. This US paper referred to our work as well.

This last condition is feasible only if there is zero liquid oxygen at take off. References 2,5,6 and 7 cited above indicate that the other two conditions are met only if air collection is done in high speed region of ascent to orbit, in our case Mach 3.5 to Mach 8. This finding was in close agreement with the first Study in 1964. We also have extrapolated our design, to show that the 1964 concept from the US can deliver 30% payload fraction only if its take-off weight is in the neighborhood of 1000-tonnes.

I now take the liberty of explaining the reason why I needed to speak to you at Salt Lake City that fateful day in June 2001. I wanted to refer you to the Paper by Jones RA and Donaldson C duR From Earth to Orbit in a Single Stage” ,from Aerospace America, August 1987, Pages 32-34.

The authors bring out here that propulsive efficiency, dry weight, and thrust-to-drag ration dominate the trade-offs in aerospaceplane design. Their work clearly bring out that it is mathematically impossible to attain orbital velocity of 8 kms per sec directly in a single stage unless the propulsive efficiency exceeds 40%, thrust-to-drag ratio exceeds 3.0, and the dry weight at take-off exceeds 56% of take-off weight.

5.Gopalaswami R “The Avatar Mission: True Pathfinder from Mini-aerospaceplane to Multi-mission Hypersonic Mass Transportation” Proceedings of International Symposium on Futuristic Aircraft Technologies, Aeronautical Society of India, Bangalore, India, 03-05 December 1996
6. Gopalaswami R et al “Spaceplanes with Aerobic Propulsion – Key to Low Cost Access to Space”, AIAA 2001-3699, 37th, AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, July 8-11, 2001. Payload in the neighborhood of 5 % of take-off weight placed in orbit with a 25-tonne aerospacevehicle carrying 56% liquid hydrogen and no lox at take-off

1.Avery WH and Dugger GL “Hypersonic Airbreathing Propulsion” J. Aeronautics & Astronautics, June 1964, Applied Physics Laboratory, John Hopkins University, Page 45, Col 2. “….design goals indicate that payloads in the neighborhood of 30% of take-off weight can be placed in orbit…”
2.Gopalaswami R et al “Concept Definition & Design of a Single Stage-to-Orbit-Vehicle Hyperplane” International Astronautical Federation Paper IAF-88-194 8-10 October 1988. Payload in the neighborhood of 15% of take-off weight placed in orbit with a 100-tonne aerospacevehicle carrying 56% liquid hydrogen and no lox at take-off
3.Balepin V Vand Tjurikov EV “Integrated Air Separation and Propulsion System for Aerospace Plane with Atmospheric Air Collection”, SAE Technical Series Paper 920974, SAE Aerospace Atlantic, Dayton, Ohio, April 7-10, 1992
4.Vandenkerchove J and Czysz P “ SSTO Performance Assessment with In-flight LOX Collection”, Acta Astronautica Volume 37, 1995, Pages 167-178,

Am I having the privilege of responding to Mr. Andrews, of “ACES” fame? Your letter requires a long answer with all the facts, so I have to split it up into several postings.

I am the person you would have (understandably) forgotten who tried to speak to you, but did not succeed in capturing your attention, even when I came up to you immediately after you presented your paper at Salt Lake City for the 37th, AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, July 8-11, 2001,
“AIAA-2001-3702 Air Collection and Enrichment System (ACES) for advanced 2nd generation RLV's“

You may now recall that I had also just presented a paper, AIAA 2001-3699, “Spaceplanes with Aerobic Propulsion – Key to Low Cost Access to Space”.

Here are the references you seek on high payload efficiency aerospacevehicles:

Space solar power as a global anti-poverty, pro-prosperity mission is a good idea.

Several attempts were made in this direction in 1990’s. These initiatives, including a private sector attempt to set up an US-Japan-India venture, did not take-off because of restrictions imposed by the US Government on transfer of information and technology in dual-use areas. I heard that even Burt Rutan faced this problem even when Virgin Atlantic a British company was to be their Spaceship II venture as a commercial partner.

Many countries now reflect US restrictions on dual-use technology. So it’s unlikely any international SSP venture can happen unless the proposed UN sponsored Outer Space Treaty to outlaw space weapons is put in place with provision for such internationally safeguarded ventures. Partner countries can then make laws to operate within this safeguarded framework.

A Correction to my posting may please be noted.
For:
a)Propulsion systems that increased the mission average fuel efficiency, and
b)For increasing the vehicle’s mission average mass ratio by extracting as much as 65% of vehicle mass while in atmospheric flight.

Please read that part as :
(a)Propulsion systems that substantially increased the mission average fuel efficiency, and
(b) By extracting up to 65% of vehicle mass while in atmospheric flight the overall mission average mission mass ratio is altered.
Put together, the product of (a) x natural logarithm of (b) the orbital velocity. This is higher and such aerospacevehicles deliver payload efficiency 15% to 30% according to later studies. This is as much as ten-to-twenty times higher than the proven Shuttle and 3 to 6 times higher than Dr. Koelle's Neptune rocket design
I’m not a mathematics man, just an aerospace engineer. The inconvenience cause by the error is regretted!!