Similarly, does suborbital release of the upper stage augment payload to orbit?
Which I am afraid is not on our website. The figures now are SI 4560 N S/kg EOL mass 1500 kg max fuel load 7500 kg with a GTO payload of 5,309 kg in reusable mode and 7,750 kg in expendable mode.
Other question: What about transferring heat from the heated (post-precooler) helium to the high pressure (pre-turbine) hydrogen via another heat exchanger? The idea is to allow far higher hydrogen outlet temps, and thus allow less of it to be used.
I've been following this topic and joined to comment on it. I'm new to this forum but have followed RE since their article in Spaceflight in 1989.Ramjets. In one of their kickoff papers RE make it clear that following the review of HOTOL they decided that any propulsion system would have to be testable on "open" facilities. Jet engines can start from rest and don't need a forced air blower in front of them. So no ramjets, pulsejets or anything close to them.I'd hoped to reference the paper but the REL website is up the spout.
And there are always the standard SSTO cautions, notably: small increases in mass can eat all your payload. With Skylon, the TPS/airframe area is so large that any last minute need to redesign it to increase strength or add thermal capacity could leave it unable to reach orbital velocity.
Quote from: Hempsell on 07/06/2011 03:26 pmWhich I am afraid is not on our website. The figures now are SI 4560 N S/kg EOL mass 1500 kg max fuel load 7500 kg with a GTO payload of 5,309 kg in reusable mode and 7,750 kg in expendable mode.That's with 1500m/s or 1800m/s of delta-v deficit? BTW, what would be the maximum payload size?
Quote from: tnphysics on 07/03/2011 04:05 amOther question: What about transferring heat from the heated (post-precooler) helium to the high pressure (pre-turbine) hydrogen via another heat exchanger? The idea is to allow far higher hydrogen outlet temps, and thus allow less of it to be used.I am afraid (like your previous proposal) without diagrams I am not quite sure what you mean. The hydrogen does not drive the main compressor turbine (it did in the old HOTOL RB545 engine). It does drive its own (i.e. the hydrogen pump) and the He circulation pump; an arrangement that means the engine can be started (which we judged to be an important feature of a operational engine). But this is already after we have exchanged heat with the helium.
Quote from: john smith 19 on 07/04/2011 08:49 amI've been following this topic and joined to comment on it. I'm new to this forum but have followed RE since their article in Spaceflight in 1989.Ramjets. In one of their kickoff papers RE make it clear that following the review of HOTOL they decided that any propulsion system would have to be testable on "open" facilities. Jet engines can start from rest and don't need a forced air blower in front of them. So no ramjets, pulsejets or anything close to them.I'd hoped to reference the paper but the REL website is up the spout. I think the paper you are looking for is Richard Varvill and Alan Bond “A Comparison of Propulsion Concepts for SSTO Reuseable Launchers”, Journal of the British Interplanetary Society (JBIS),Volume 56, pp. 108-117, 2003It is on our website Media Library – pdf downloads and it was working just before I posted this. I am sorry if it gave you trouble.Of course, we do have a bypass ramjet and this cannot be verified by static testing alone, but we are being pessimistic in its performance estimates and we will have a flying test bed to ensure we have it right.
I think the paper you are looking for is Richard Varvill and Alan Bond “A Comparison of Propulsion Concepts for SSTO Reuseable Launchers”, Journal of the British Interplanetary Society (JBIS),Volume 56, pp. 108-117, 2003It is on our website Media Library – pdf downloads and it was working just before I posted this. I am sorry if it gave you trouble.Of course, we do have a bypass ramjet and this cannot be verified by static testing alone, but we are being pessimistic in its performance estimates and we will have a flying test bed to ensure we have it right.
Okay. What is the schedule for SKYLON? How likely is it to be built?
I know that amongst the British team, they're committed to the ideal of a one-body flight vessel, but the three main things I can see that would really get the payload fraction up (without compromising the use of commercial runways, rapid turn-around of flights and remarkable economy) are1. JP-7 "traction shell" -- a la SpaceX, a conventional titanium wing and SR-71 Blackbird engines to cover the ground-to-60,000 feet at Mach 6. There are dozens mothballed, and are magnificent billion-dollar engineering marvels. Use 'em. They also cover both turbo and ram-jet operational modes. Use 'em. When the payload (the Skylon) is at altitude... fire its engines, and peel away. The traction pilot lands conventionally.2. Solid rocket hypersonic insertion -- Next, fire up a set of SRBs. Push the bird to a steep angle of ascent, and get to Mach 12. Drop the SRBs. SRBs have winglets and return "almost" to base, parachuting the last 5000 feet into a body of water. Saves a lot of money recovering by Navy way out at sea.3. Hydroplane landing -- While this challenges the "commercial runway" idea (which is a farce anyway), few spaceports couldn't be outfitted with a hydroplane landing strip. 6 inches of water, 100 meters wide, 2,500 meters long. The landing gear are just pontoons of a sort - without even the need to be hollow as there isn't enough water to sink into. Curved hydrodynamically to spray water laterally, there becomes no need for heavy landing gear. Skid-stop. These answer to a lot of technical problems that the British team has engineered sub-optimal ways around. Taking off with tons of water (for cooling the brakes in an emergency take-off abort), and having landing gear that is strong (heavy) enough for the whole mission is just a big compromise. The skids shouldn't be more than 10% of the weight of landing gear. All that mass becomes payload.Likewise, 100% of the hydrogen NOT consumed getting from zero to Mach 6 becomes payload. The traction craft has conventional "big tire" landing gear, can effect a safe abort, uses MUCH less expensive JP-7, utilizes already-engineered ramjet engines, and improves the economy markedly. Even in its special formulation, JP-7 is 1/20TH the cost of liquid hydrogen for 1 unit of energy.Finally the use of SRBs again serves the purpose of taking the majority of the "dead weight" (unburned fuel) and gives it a big old kick in the butt without needing to burn so much of it.The same SKYLON craft very well could make it on its fuel all the way to MEO or even Geosynchronous. Or, with all that tankage - it could just transfer all its excess hydrogen and oxygen to orbital tanks for other missions to use for satellite transfer and the like. There is ALWAYS use for "extra" hydrogen and oxygen in orbit.The use of "conventional" runways is undermined by (a) the teams acknowledged high-loading on tiny landing wheels, and (b) by the need for special hydrogen storage, delivery, firefighting and substance control at the airport. No, the "conventional" side is just that a military-grade or aerospace port would be required, with all its attendant specializations. This is not to say that a NEW airport would need to be made, no. Just special. [...]The traction shell is really a formal first stage. I don't imagine there is a compelling reason to burn both the JP7 and the hydrogen (first and second stages ganged) since the winged traction shell would gain altitude using aerodynamic lift instead of raw impulse. Might just as well save the hydrogen, and use a smaller (lighter) tank.I don't think there is a downside for having wings on the second (orbital) stage both for rarefied-atmosphere lift, and for conventional runway recovery. They do add weight, but with the traction shell is going to out-weigh the orbital stage probably by 2× to 3× It can afford to be a muscular meatball - aerodynamics is going to be the lever that converts its thrust to potential (altitude) energy.Most "rocket scientists" propose getting as high as possible quickly - because that is the 60+ year convention of rocket science, not because the alternate lifting-body dynamics makes substantially less sense. Even in the 1950s, rocketry professionals. were considering aerodynamic body dual-mode systems such as I advocate. There were two main problems which had no technological answer: engines that were efficient enough to run in the 50,000-to-200,000 ft stratosphere at super-to-hypersonic speed, and airframe materials simultaneously light-weight, thermally stable, strong and inert enough to serve aeroframe duty under the extraordinary thermal, pressure and flexure regime of "near space".That picture has now changed. The magnificent work alone by the Japanese through the 1980s-to-2000 (and not over by any means!) in creating outrageously durable, tough, resilient refractory ceramics is the key to the "leading edge" problem both for high-atmosphere aerodynamic loading on launch, and the very-high atmosphere compression on reëntry. These same ceramics make possible cowlings and compression cones for the hybrid ramjet/scramjet engines that in turn power the whole thing well into the mesosphere (200,000 to 500,000 ft or 60-150 km).So I claim that it is the convergence of exemplary materials properties and a new generation of "rocket scientists" that can now explore such unconventional designs. Ultimately, we are heading toward space ports of the type only imagined by science-fiction writers of the 1950s.[...]RAM/SCRAM jet engines suffer from the very riches that make them such compelling designs. The faster the inlet airstream, the higher the compression that can be achieved within the speed-of-sound cone. The problem is [ PV = nrT ] ... which implies that as pressure goes up (and volume goes down), that temperature will inevitably need to rise (in an adiabatic system). The gasses are rushing into the compression system well above the speed of sound, so the compression doesn't have any time to become non-adiabatic (i.e. to lose heat).The current strategy - as I understand it - is to try to cool the airflow using heat exchangers, which given the speed of gas-flow, seems hopelessly optimistic. Maybe its not - after all, the Shuttle's (and virtually every rocket ever made) uses the cryogenic oxygen and/or hydrogen to cool the expansion cone and interior of the engines (along with a "film" of unburned fuel). The one way by which it could work out well for the hybrid engine would be if the amount of liquid oxygen is increased, and it is sprayed directly into the compressed inflowing air. The enrichment wouldn't hurt the hydrogen burn, and it would definitely cool the stream.In any case, one of the most compelling reasons for going "two stage" with a traction lifter is that the kind of engine can be optimized for its operating regime. The orbital stage can have its no-spinning-parts SCRAMJET engine (which is magnificent, and highly efficient), which would be much lighter weight, and much more free from the problems of supersonic compression and heat dissipation. Indeed - I'm unable to muster a single compelling reason to try to make a single engine perform both duties.