Well, I thought maybe I could add some content. Everybody’s heard about the British Hotol follower Skylon and its airbreathing SABRE engines.
What’s special about them? My understandin’s based on this excellent document from Reaction Engines explaining why the system is what it is.
What they are and what they are not
Well, they are air breathing engines that run on hydrogen. They precool the intake air with hydrogen and have an intake air compressor like a jet engine so they can start from a standstill.
They are not ramjets (and definitely no scramjets since the flow is subsonic) since they have a compressor and thus also have better compression ratios and efficiencies, meaning less airflow needed for the same thrust.
They are not that far from air turborockets, but they do a neat trick, and they are very close to LACE – liquid air combustion engines – except that they don’t liquify the air.
The basis – RB545 of Hotol
SABRE is conceptually based on the planned Rolls Royce RB-545 powerplant of the Hotol – an aborted British spaceplane of the eighties. The RB545 was to use hydrogen to precool the air before compressing it with a compressor. Then some of the now hot hydrogen was used to drive a turbine (that would drive the compressor and the hydrogen pump) and exhausted while some of it was fed into the main combustion chamber together with the high pressure air.
This concept was quite close to a LACE engine where the incoming air is cooled all the way until it liquifies. Then it requires very little pumping effort (pump size and power is related to volume flow, not mass flow). You can also separate the oxygen and nitrogen in this way.
On the other hand it resembles a turbojet since it has a compressor, or an expander air turborocket (but there the cooling is not used beneficially).
The problem with these kind of engines is that they require an excess of hydrogen for the cooling. Now, RB545 and SABRE are less extreme than LACE engines – they only cool the air and don’t liquify it.
In LACE, only a small part of the hot hydrogen coming from the intake heat exchanger is used for work, for compression of oxygen and hydrogen. They are liquids anyway and require little effort to pump. So a lot of hydrogen needs to be carried but most of it is then left unused after it’s been used to liquify the intake air.
On the other hand, in a turbojet or air turborocket, the air (especially at high speed and altitude) is warm, fluffy and requires a huge heavy and draggy compressor to compress and that requires a lot of power too.
It turns out that a good middle ground can be reached where the supplies and the needs meet. Only cool the incoming air enough that you can extract the compression effort needed for it from the hot hydrogen you just heated with the air. This is the whole idea of RB545 and SABRE.
Now, you still need slightly more hydrogen for this than you’d really want. Rocket engines run in the 5.5x to 6x LOX-heavy category (and they’re still rich, because OH would be 1 hydrogen to 16 oxygen masses and H2O 8:1). RB545 was designed to go with a 10x air mass compared to hydrogen mass. At a 20% oxygen concentration, that’s only 2x oxygen mass. SABRE is designed to be at 12.5x air or 2.5x oxygen.
The cooled air is nice since now the compressor can be smaller than in a jet with a similar mass flow. Also it can be made of low temperature materials. The inlet can be much much simpler and lighter as it doesn’t need to do so much compression as in a hot engine. This is important as inlets usually dwarf engine masses in large Mach range airbreather designs.
SABRE doesn’t really use hydrogen in the precooler and doesn’t exhaust excess, instead it has a secondary Brayton cycle helium loop. Apparently this is beneficial for turbine/compressor matching and avoiding hydrogen embrittlement. The rocket chamber also needs cooling and this is planned to be done with liquid oxygen. There are of course hydrogen, LOX and helium pumps, a hydrogen preburner and a few heat exchangers and many other things left out of the simplified diagram.
What does it mean?
I see four large weaknesses in Skylon.
- The precooler might freeze from air moisture. If the LOX troubles by various rocket teams are anything to go by, nevermind hard cryogens… This is a very central known issue about the design.
- The craft needs a lot of hydrogen, resulting in a very awkward structure. How about flying a Zeppelin at Mach 5? And remember that it has to use a lower trajectory than a rocket resulting in aero loads. Also an acknowledged thing.
- It’s still a low margin SSTO where the payload at 4% of total weight is very sensitive to performance downgrades in the rest of the vehicle. The whole study is based on the premise of SSTO. What if you stage instead? Won’t staging rockets eat Skylon’s lunch?
- When the oxygen mass used is only 2.5x the hydrogen mass carried, are you saving much by airbreathing? If you look at the volume of the vehicle dedicated to airbreathing hydrogen carrying, you could use 6/7 of it for hydrogen and 1/7 of it for LOX – with the 20x density of LOX it would hold 3x the original hydrogen mass of LOX. 3x > 2.5x – it would carry more oxygen than the Skylon uses in its whole airbreathing operation! Of course we must keep in mind that the engine is thrifty and the nitrogen reaction mass going through the engine is useful as well.
I’m still undecided with Skylon. A SABRE related engine might make for a good first stage airbreather anyway. I don’t know if it would be of any use to run any similarly concepted engine with methane.
Apparently this is beneficial for turbine/compressor matching and avoiding hydrogen embrittlement.
It also avoids patent issues, something they mentioned in early papers, but don’t stress anymore…
1. The Japanese solved that problem in 1988 when I saw Mitsubishi’s 1 cubic meter LACE heat exchanger run. If you take Balepin/Rudakov they were adament deeply cooled advocates. Leingang at Wright-Patterson was just as adament for LACE. In the end its six of this half-dozen of that. Very little weight difference.
2. Hindenberg was gaseous. Whether an airbreather, combined cycle of rocket the energy to orbit is about the same as is the mass and volume of hydrogen. Its the LOX that is heavy and you want to remove.
3. TSTO is the way to go regardless of the propulsion. The key is dry weight payload fraction. The gross weight is just operational empty weight time mass ratio.
4. If I have 100,000 lbs of H2 it takes 600,000 lb of LOX to use a rocket. If I have a combined cycle it only takes 250,000 of LOX, a 350,000 lb saving!
1. I’d love the hear more from this stuff! I don’t know the subject matter very well, but the Skylon guys’ paper linked makes a convincing case against LACE. Like Ian said, too much hydrogen needed.
2. True, that was hyperbole. At 70 kg/m^3 liquid hydrogen is 7% the density of water but still 60x as dense as air. I think some high strength sandwich foams like Airex are in a similar category.
It’s still very bulky compared to LOX at 1100 kg/m^3. And the aero loads are in a different class than on a Zeppelin.
4. But that LOX takes so little space (meaning tanks are light and cheap) and rocket thrust is easy as well (light and uncomplicated). Does it matter in the end? You need inlets, compressors, heat exchangers, wings, an awkward trajectory and different flight modes with their aerodynamic stability problems and structural loads… lots of things just to save that dimensionally small (although heavy) amount of LOX. Which will have more *dry* mass in the end, the one with a lox tank and more rocket engines or the one with wings, inlets, compressor, heat exchanger, support frame? And more importantly, which will be cheaper?
I’m saying, it’s not clear cut.
You say that 4% is low, but it’s still higher than the staging, Shuttle design, or even the staging, non reusable rockets, and is much higher than any pure rocket SSTO approach is ever likely to be.
They’ve already tested the precooler, and shown that it doesn’t freeze up. They’re able to condense the water vapour, and expel it in liquid form before it freezes.
LACE simply doesn’t look as good as this, it uses too much hydrogen, and you end up dumping the hydrogen through a propulsive nozzle, giving pretty good ISP (about 800 seconds), but it doesn’t compare with the 2000 seconds from SABRE; and 800 seconds isn’t quite enough for SSTO.
I think a better term would be payload sensitivity to stage dry mass growth. In a two stage rocket this sensitivity can be an order of magnitude lower than in a single stage rocket. I’ll post on it.
I’m sure that they could go two stage if they *had* to, and then they would have really *massive* payload; because the aeroengine portion seems to be a really big win, I think they’re roughly quadrupling the payload with that engine scheme (but I’d have to check my maths on that.)
There would be some issues with the release though, releasing at 26km/Mach 5 is unwise, they would probably have to do a zoom climb for a 2 stage system.