Spacex has been awarded a $40,766,512 modification (P00007) for the development of the Raptor rocket propulsion system prototype for the Evolved Expendable Launch Vehicle program. Work will be performed at NASA Stennis Space Center, Mississippi; Hawthorne, California; McGregor, Texas; and Los Angeles Air Force Base, California; and is expected to be complete by April 30, 2018. Fiscal 2017 research, development, test and evaluation funds in the amount of $40,766,512 are being obligated at the time of award. The Launch Systems Enterprise Directorate, Space and Missile Systems Center, Los Angeles AFB, California, is the contracting activity (FA8811-16-9-0001).
SpaceX has conducted several dozen successful hot fires of the Raptor engine. SpaceX has completed over 1,200 seconds of firing across 42 main Raptor engine tests.
At the IAC meetings September 2017, Elon Musk announced that a smaller Raptor engine — with slightly over half as much thrust as the 2016 proposed designs — would be used on the BFR rocket than had been used on the ITS launch vehicle design unveiled a year earlier. Additionally, fewer engines would be used on each stage. BFR would have 31 Raptors on the first stage and 6 on the second stage, whereas the ITS launch vehicle design had 42 larger Raptor engines on the first stage and 9 of that same large size on the second stage. The engines remain full-flow staged combustion cycle design using subcooled liquid-methane/liquid-oxygen propellant, just like the 2016 larger engine designs. “Version 1” of the flight engine is designed to operate at 250 bar of chamber pressure; but SpaceX expects to achieve 300 bar in later iterations. The flight engine is designed for extreme reliability, aiming to support the airline-level of safety required by the point-to-point Earth transportation market.
The sea-level model Raptor engine design, with a nozzle exit diameter of 1.3 m (4.3 ft), is expected to have 1,700 kilonewtons (380,000 lbf) thrust at sea-level with an Isp of 330 s increasing to an Isp of 356 s in the vacuum of space.[48] The vacuum model Raptor, with a nozzle exit diameter of 2.4 m (7.9 ft), is expected to exert 1,900 kN (430,000 lbf) force with an Isp of 375.
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It’s odd that the USAF would give SpaceX more money immediately after SpaceX had so publicly bet the farm on Raptor’s development, especially when the original OTA specified only engine development. Somebody on another thread suggested that the extra $40M may actually be the USAF pulling the pin on an undisclosed option to have SpaceX actually develop a prototype Raptor upper stage (RUS) for the F9 and FH.
After doing a bit of figuring (which I’ll spare you), I came up with two surprising observations:
1) If you subcool methane to near slush (density = 450 kg/m³), and if Elon’s stated O/F mixture ratio of 3.8 is actually correct (that’s incredibly O2-rich–optimal at 250 bar is about 2.95), then you only need 6% more volume to produce a methalox RUS with the same energy as the existing kerolox F9 S2. That makes the RUS pretty easy to develop as a 3.66 m diameter stage, with less than an extra meter of vertical tankage.
2) If you assume about a 25% CAGR in launches for SpaceX, about 20 block V cores with a reusability of 50 launches each, and roughly current production levels, you can build all the cores and Merlins you need to launch through 2025 by the end of 2018. But to build all the Merlins for the expendable second stages would take you until 2022 to develop.
This seems like a really good biz case for a RUS. If you can effectively shut down the Merlin line at the end of next year and start ramping Raptor production to build RUSes, you get lots of early manufacturing experience with the Raptor and free up the resources necessary to make the BFS and BFB test articles for the BFR system.
Found the answer myself the BE-3 is a hydrogen engine and the BE-4 is methane.
Isn’t Blue origins engine using hydrogen not methane
“But LH has so much more energy per kilogram”
Yeah, pity you can’t burn it without all that oxygen.
Yep. Too bad one couldn’t deploy a thousand tons of mercury (in a TINY volume compared to conventional cryogenic materials). And ions. And beamed power. Maybe 5 GW worth. 500 kg/s up front, then throttle down the kg/s, increase Isp as mass drops.
Its an interesting integration problem, actually. Not the hardest, but certainly has its challenges.
GoatGuy
PS: consider using the reply button?
How about beamed power with liquid nitrogen fuel? N2’s Triple bond will soak up the Joules.
From this handy document, Table 1 on page 2 Liquid Nitrogen has an ISP of 253 at 2800 degrees kelvin.
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910012833.pdf
Might want to switch to water which has an ISP of 370 at the same temperature.
Beamed power in the multiple GW range would make for some very simple rockets-propellant combinations for SSTO rockets.
If you’re going to do “beamed power” better to go with hydrogen (isp 1,000!).
I will when I can; At work I have to use an obsolete browser, and some of the site’s functionality is missing. “Reply” is one of the missing features.
Interesting possibilities. It would appear that the Raptor is about the same size or maybe slightly larger than the Merlin. I’m no math major, but it seem to me that if you built a “Falcon 5” using 5 Raptors, you would have a very slight increase in thrust, a 6-7% increase in ISP and engine weight savings of over two tons compared to the Falcon 9 FT to offset any weight gain from having two cryo fuel tanks. I have read in other places that the USAF is quite interested in a Falcon 9 that uses the Raptor vacuum for the second stage. My understanding is that increased impulse is more important on the vacuum stages than on the booster stage so that might result in an appreciable increase in payload capacity, especially to GTO/GSO
I am a bit of a math nut, as you might have surmised in reading NBF. here’s the analogy that works: if I give you a hand-held rocket that produces 9.8 N and weighs 1 kg, and doesn’t get lighter during firing (or for that matter doesn’t burn out) … and you point the business-end toward earth (so that it might fly away)… will it? Nope. Within precsion, that 1 kg is attracted to earth with 9.8 newtons of force. The rocket will be pushing the other direction at the same force. When you let the rocket go, it’ll just float around indefinitely.
This is of course NOT how real-rockets work. But it does illustrate: you’ll burn a lot of rocket fuel for nothing. Overcoming gravity.
REAL rockets however get lighter as they whizz away. They start off heavy. And to fly away, they have to have more thrust than the gravity attraction. They rise up, getting lighter, the thrust nearly constant (it can be throttled of course). So upward velocity starts to accumulate. And they get lighter, so have even greater thrust going to acceleration, so go much, much faster. At some point they’re going fast enough to break the speed-of-sound barrier; the pushing-away of the atmosphere in the front again acts to retard their acceleration. But they’re also getting lighter. And so it goes… Once empty-ish, the first stage is shucked as being “useless weight”, leaving the second stage to fire up and get going. And so on…
Your thoughts are on the right path tho.
Yes – once you are essentially free of climbing out of the Earth’s gravity well into airless space, then to conserve fuel you prefer higher ISP than highest thrust. Remember that highest thrust offsets the “dead weight” of the full stack at liftoff.
In fact, ideally, the first stage (as has become something of the norm these days) would have EXTRA burners, that would fire for maybe just a few minutes at a stupendous rate … to maximally overcome gravity and expend most of the energy toward gaining upward and forward velocity.
That’s the essence of rocket science.
GoatGuy
All the while limiting max Q. Too much acceleration early on and you incur too much pressure when low in the atmosphere. F9 and other rockets throttle down before max Q so that they don’t rip themselves apart due to the pressure stress on the frame.
Rocket optimization math is quite involved. Fortunately it was mastered some time ago.
I do love beamed power though. I wonder if instead of using H2 for fuel with beamed power you could use liquid nitrogen without getting too much high atmospheric smog. Nitrogen’s triple bond strikes me as being able to soak up the Joules and being heavier than H2 it should provide better thrust.
“Nitrogen’s triple bond strikes me as being able to soak up the Joules and being heavier than H2 it should provide better thrust.”
But you don’t want to “soak up the joules” do you? You want something that have very low specific heat so that you get max exhaust temperature per joule transmitted by your power beam. Helium or something is best for that.
But then you want an atom that is super light, because that gives you the max exhaust velocity for the given exhaust temperature. Temp being limited by the rocket chamber and nozzle materials that have to withstand the temp without failure. So hydrogen is best there. With helium at second place.
But then you want something that is super compact and doesn’t need cryogenic tanks. Because those tanks are heavy too. And huge volumes of liquid hydrogen, as you say, take big heavy insulated tanks. Liquid nitrogen is better, but not ideal.
So you want something that forms a dense liquid when at room temperature so you can keep it in a small, cheap tank and then feed it to the motor.
You might end up with hydrocarbons after all, even with beamed power. Only just using them as reaction mass without the need for oxygen.
The cheapest reaction mass is clearly water.
Ignoring the cost of the reaction mass, lithium? Density as a liquid is nearly half that of water, the atom is super light, it melts at only 180 which is probably OK for some aluminium alloys, and certainly everything else. You’d have to preheat it before launch…
Hydrogen is best for ISP but it isn’t cheap and it is hard to handle. Water would be great. Non cryogenic.
The double bond of nitrogen doesn’t break easily so the propellant isn’t corrosive to the rocket nozzle or heat exchanger. (ISP is low though).
We may see a day where propellant choice and cost is a determining factor.
The basic compromise between going tangentially (horizontally if you’re in flat-earth mode, which is a good approximation for the time you’re in the atmosphere), and radially (vertically) is the “gravity turn”:
1) Go straight up, at as high an acceleration as possible, until you hit some fairly small critical velocity.
2) Execute a “gravity kick”, which is a very small angle (often less than 1°) pitchover.
3) After that, guidance simply aligns the thrust vector with the velocity vector. Because there’s a downward component to the velocity, the vehicle will gradually curve over to tangential thrust, but do it high enough up not do incur bad dynamic pressure. The other nice advantage of this is that, except for the kick maneuver, you’re always at 0° angle of attack.
4) Gravity turns stop being as useful once you’re nearly in vacuum, and other methods are used to shape orbital insertion properly.
The math isn’t so much “mastered” as it is solved with a grab-bag of numerical tools. There’s still a modicum of black magic involved.
The problem with nitrogen is that it’s heavy, and specific impulse (which is really just exit velocity / 1 gee) is maximized for the lightest particles you can get. Hydrolox comes out as a thermodynamically unpleasant combination of H2O, H2, O2, H-, O+, OH, and so on, but the average molecular weight is about 13.5–in other words, it’s got lots of monatomic O and H. The very, very best you can do with N2 is 14. Best propellants for beamed power are H2O, CH4, AlH3, and H2.
You don’t want to maximally overcome gravity; you want to get going horizontally (tangentially) as soon as possible. After that, high enough speeds generate centripetal acceleration, and you’ll naturally achieve the altitude you want.
But you can’t go fast horizontally until you deal with the atmosphere generating too much dynamic pressure. Hence the “gravity turn”, which is a nice, simple compromise between getting high fast (and therefore out of the atmosphere) and minimizing gravity losses (where you want to get horizontal).
So, get just high enough, and get horizontal. In the immortal words of Slim Pickens, “Shoot, a fella’ could have a pretty good weekend in Vegas with all that stuff.”
Correction to my previous to GG: Centrifugal acceleration. Inertial frames are for wimps. Embrace the rotation!
Putting 5 Raptor engines on a Falcon-9-sized rocket would make it an expendable-only booster with no way to recover it. No way you can throttle down a raptor engine low enough to allow a retropropulsive landing– The Raptor engine is way too powerful for that.
One of the biggest reasons why the Falcon 9 works as well as it does (especially for landing the booster so it can be reused again) is the fact that one single Merlin engine at its minimum throttlable 50% power is just enough thrust to land the nearly-empty booster with a reasonably gentle hoverslam.
Pretty shock cones, there. Purple no less. Cool.
I’m reminded about the almost ridiculous condition that exists in the Space Shuttle main (not booster) engines. Where the amount of expansion is so great and the temperature of the expansion cone is so low that liquid water condenses onto it in operation. You can see it streaming out of the outside edge in some really close-up STS launch videos. Incredible!
Methane is 2.5× more energy-dense (per liter) than liquid hydrogen. But LH has so much more energy per kilogram. And … once you get more-or-less out of the atmosphere, it is energy density PER KG that counts.
GoatGuy
Not so fast…
Less dense fuels like hydrogen need larger tanks which means more mass.
Furthermore liquid hydrogen needs insulation, i.e. even more mass.
So it’s not just the density of the fuel that matters.
All things considered I think a vehicle with a given payload mass to orbit will be much larger (so more expensive) if it uses hydrogen as opposed to dense fuels (RP-1, CH4 etc).
“Methane is 2.5× more energy-dense (per liter) than liquid hydrogen. But LH has so much more energy per kilogram. And … once you get more-or-less out of the atmosphere, it is energy density PER KG that counts. ”
While that is true, the energy density is *not* the only thing that matters in a fuel, even beyond the atmosphere. Availability counts! Since BFS can get Methane made on Mars, it is designed to run both its stages on Methane. In addition, Hydrogen has silly little tricks like leaking 100 times more than Methane, and like Hydrogen embrittlement of the tanks and especially the plumbing in rocket engines that shake at high gee forces whenever running. If water is found on the Moon, whether in polar craters or in lava tube caves, then Hydrogen will be an excellent propellant for the New Sheperd landers on the Moon that Blue Origin plans to use there. Ultimate use for a propellant counts heavily in selection because of availability. Even the Mercury you mentioned might one day come into use as a propellant, once ion engines using it can be tested in high Earth Orbit where the Mercury won’t contaminate humans or their environment. Then we carefully build the engines and vehicle they power at EML-1, move the vehicle out of plane for the factory they are built in, and load Mercury propellant, turn on the power source, and enjoy the highest thrust density we can get out of ion engines we can calculate.