3 ms·

It's one more tool for achieving reuseability without sacrificing performance. The size and weight of an air-breathing system will be lower due to greatly incr

It's one more tool for achieving reuseability without sacrificing performance. The size and weight of an air-breathing system will be lower due to greatly increased Isp of the 1st stage, the largest stage by mass. A smaller 1st stage fuel tank might mean that heat-shielding or higher temp materials is more of an option, which means you can raise the 1st stage separation velocity closer to what is "optimum" in the rocket sense and still recover the 1st stage. This puts the 2nd stage closer to its natural optimum design point. Overall size might not be a show-stopping problem now, but for building space colonies or Mars missions, overall vehicle size does eventually become concerning. Think of trying to build a jig to friction-stir weld a 180m long, 14m wide fuel tank. It's cool but also a bit of a distraction to have to build and handle everything at unusual scale. It would be super cool to see big packs of ten F414's at the base separate off, curve back towards the launch site, flip out some wings, and then cobra into a vertical landing. GE should step in right now to have the perfect platform to develop ceramic parts without the same service life requirements as aviation gas turbines.

Gravityloss 12y ago
Cool, yes, but the cost effectiveness might be questionable. Rocket engines produce ten times as much thrust per kg of hardware. If you're doing it well, they should be relatively cheap and maintainable. LOX tanks are also small since LOX is so dense. I wouldn't build a 180 m long 14 m wide fuel tank. I'd launch lots of smaller missions and do space refueling, assembly or docking. None of the components of any mission are really credibly over 20 tons dry. Look what the air forces do - they don't build massive aircraft with excessive range - they do aerial refueling. Space launch is an acceleration mission. Your goal is over 8 km/s. The air breathing turbine engines can help perhaps a few hundred meters per second, for a lot of dry weight, and that's nice. With wider speed ranges they get even more complex with finicky inlets and produce diminishing results. There's been some cool engines in the history that could provide some interesting capabilities http://en.wikipedia.org/wiki/General_Electric_GE4 http://en.wikipedia.org/wiki/General_Electric_GE4 I don't know what a jet engine designed for acceleration would look like. I think the most potential in air breathing lies in air launch of relatively small rockets. Dropping the rocket from a large subsonic airplane. It brings a lot of flexibility with launch sites and times, plus abort options for the dropped vehicle.
- knappador 12y ago
  I did some back-of-napkin calculations on the Falcon 9's 1st stage numbers here: http://www.spaceflight101.com/falcon-9-v11.html http://www.spaceflight101.com/falcon-9-v11.html Including the rest of the dry mass fraction of the vehicle, (the tank, the structure to hold the tank etc), each Merlin's dry thrust-to-weight ratio drops to 43.05. (654kN * 1000 N / kN / 9.8 (kg * m/s2)) / (Inert mass / 9.0). The GE F414 in the high-thrust configuration is the dry-mass comparison using COTS. Let's say it's 8.0 when factoring in some control system and inlets. Isp of a jet engine burning liquid methane is ~2000 even in full reheat. Isp of rocket engine is ~330 for methane-lox in the newer full-feed Merlin. The impulse to fuel weight ratio is almost 10x inverse in favor of the air-breathing system. The dry mass thrust-to-weight is decisively in favor of the rocket engine. The wet mass thrust-to-weight is decisively in favor of the jet engine. Affect of these factors on cost is inconclusive from the napkin. How can air-breathing engines potentially lower the TCO? 46 reusable RS-25 were flown on 135 shuttle missions, so the average lifetime of each engine was...8. The lifetime of a gas turbine is measured in thousands of hours, so let's say 200 missions and without so many major overhauls. Even if Space X is pulling out some cool voodoo for Merlin reuse-ability, gas turbines are in a different league in terms of both reliability and lifetime. While a 1st stage rocket body has to re-enter, land, and be transported back to the launch site, an air-breathing 0.5 stage can fly itself back from its short distance downrange, thereby incurring zero transport cost back to the launch site. An extreme application of air-breathing might include ram-jet inlets that opened after mach ~1.8 and took over air-breathing duties up to Mach 4.0 as the 414's idled. Even in that case, the whole system is not that far downrange and can use altitude to make up quite a bit of the journey back and use some vertical landing system if it saves weight. The fuel efficiency lets you do this with 414's in spite of their lower T:W. Does using air-breathing assist pay off by making a lighter 1st stage or heat-shielding the 1st stage to make a faster release point? Only extensive analysis of the various options can answer these questions. I believe the mechanisms are there, and that's what justifies the design research.

Gravityloss 12y ago

Cool, yes, but the cost effectiveness might be questionable. Rocket engines produce ten times as much thrust per kg of hardware. If you're doing it well, they should be relatively cheap and maintainable. LOX tanks are also small since LOX is so dense. I wouldn't build a 180 m long 14 m wide fuel tank. I'd launch lots of smaller missions and do space refueling, assembly or docking. None of the components of any mission are really credibly over 20 tons dry. Look what the air forces do - they don't build massive aircraft with excessive range - they do aerial refueling. Space launch is an acceleration mission. Your goal is over 8 km/s. The air breathing turbine engines can help perhaps a few hundred meters per second, for a lot of dry weight, and that's nice. With wider speed ranges they get even more complex with finicky inlets and produce diminishing results. There's been some cool engines in the history that could provide some interesting capabilities http://en.wikipedia.org/wiki/General_Electric_GE4 http://en.wikipedia.org/wiki/General_Electric_GE4 I don't know what a jet engine designed for acceleration would look like. I think the most potential in air breathing lies in air launch of relatively small rockets. Dropping the rocket from a large subsonic airplane. It brings a lot of flexibility with launch sites and times, plus abort options for the dropped vehicle.

knappador 12y ago
I did some back-of-napkin calculations on the Falcon 9's 1st stage numbers here: http://www.spaceflight101.com/falcon-9-v11.html http://www.spaceflight101.com/falcon-9-v11.html Including the rest of the dry mass fraction of the vehicle, (the tank, the structure to hold the tank etc), each Merlin's dry thrust-to-weight ratio drops to 43.05. (654kN * 1000 N / kN / 9.8 (kg * m/s2)) / (Inert mass / 9.0). The GE F414 in the high-thrust configuration is the dry-mass comparison using COTS. Let's say it's 8.0 when factoring in some control system and inlets. Isp of a jet engine burning liquid methane is ~2000 even in full reheat. Isp of rocket engine is ~330 for methane-lox in the newer full-feed Merlin. The impulse to fuel weight ratio is almost 10x inverse in favor of the air-breathing system. The dry mass thrust-to-weight is decisively in favor of the rocket engine. The wet mass thrust-to-weight is decisively in favor of the jet engine. Affect of these factors on cost is inconclusive from the napkin. How can air-breathing engines potentially lower the TCO? 46 reusable RS-25 were flown on 135 shuttle missions, so the average lifetime of each engine was...8. The lifetime of a gas turbine is measured in thousands of hours, so let's say 200 missions and without so many major overhauls. Even if Space X is pulling out some cool voodoo for Merlin reuse-ability, gas turbines are in a different league in terms of both reliability and lifetime. While a 1st stage rocket body has to re-enter, land, and be transported back to the launch site, an air-breathing 0.5 stage can fly itself back from its short distance downrange, thereby incurring zero transport cost back to the launch site. An extreme application of air-breathing might include ram-jet inlets that opened after mach ~1.8 and took over air-breathing duties up to Mach 4.0 as the 414's idled. Even in that case, the whole system is not that far downrange and can use altitude to make up quite a bit of the journey back and use some vertical landing system if it saves weight. The fuel efficiency lets you do this with 414's in spite of their lower T:W. Does using air-breathing assist pay off by making a lighter 1st stage or heat-shielding the 1st stage to make a faster release point? Only extensive analysis of the various options can answer these questions. I believe the mechanisms are there, and that's what justifies the design research.