John Bucknell presented his updated Turbo Rocket plans at Space Access 2019.
John has previously worked on the SpaceX Raptor engine and has presented a nuclear air-breathing rocket design. The new design is a single-stage to Orbit air-breathing rocket.
It would have a mission-average Isp of more than 1,600sec and payload fractions of 25-50%
John revisited the assumptions around rockets and will use hot hydrogen fuel to enable air-breathing hypersonic rockets to be simplified.
An air-breathing rocket can increase the payload capacity from 1-4% to 25-50%.
The rocket would run at 1000 to 1300 kelvin which does not require new materials.
There has been previous development of hot hydrogen vehicles.
It is a lighter rocket and will use air-braking and then a propulsive thrust landing. Vertical takeoff and vertical landing.
John has a single gas path.
John now has a non-nuclear staged combustion design.
The longer-term plan would be to use this low-cost reusable launch for energy from space-based solar power.
Prometheus Medium would be able to deploy 7,000-8,000 subsats per launch in racks of 100. This array is the first 14,000 elements of an 81,000 element array for 100MW – requiring 10-12 launches. The phased array microwave transmit aperture at 1.5km is big enough to illuminate a 3.5km diameter rectenna from 35,000km apogee and can scale to 1 GW by filling in the sparse array. 5.8GHz frequency is able to beam through atmosphere, weather and darkness with no loss. Baseload with a highly eccentric Molniya orbit requires three satellites in the same orbit, 8 hours apart – costs account for 66% availability.
John has designs for a more efficient and lower cost space-based solar power than previous proposals.
It is a distributed array. There would need to be replacement of the satellite elements in the array.
SOURCES- Live coverage of John Bucknell presentation at Space Access 2019. Interview with John Bucknell.
Written By Brian Wang
Brian Wang is a Futurist Thought Leader and a popular Science blogger with 1 million readers per month. His blog Nextbigfuture.com is ranked #1 Science News Blog. It covers many disruptive technology and trends including Space, Robotics, Artificial Intelligence, Medicine, Anti-aging Biotechnology, and Nanotechnology.
Known for identifying cutting edge technologies, he is currently a Co-Founder of a startup and fundraiser for high potential early-stage companies. He is the Head of Research for Allocations for deep technology investments and an Angel Investor at Space Angels.
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A new edition has been posted on YouTube including narration and more detail: https://youtu.be/wa3eIQHmf2I
John
This presentation has been taken down, new one with narration and more detail up. See post above.
The reason I asked that is because I’m thinking of how manned flight for long terms and such would work. Compared to SpaceX’s Dragon capsule, Boeing’s capsule and the Orion capsule how is potential pressurized volume, and passenger size limits.
Single stage non-return payload is pretty small – on the order of 2-3% of LEO payload with the nuclear variant. Refuelling in LEO (one additional launch) gets you 50% payload and return to terra firma. These values are for a 5GWth version of the rocket – 11m core diameter, six times the payload of the 6.5m core 1GWth initial vehicle.
I have a DVD of Destination Moon and I showed it at Starship Congress in 2017. The accuracy considering the time frame was really quite good – with the exception of having a 3000s Isp nuclear rocket.
Shielding from gamma rays is basically mass between you and the sources, about 5km of air is more than enough – about 50mm of tungsten is equivalent. Neutrons need a hydrogen atom to run into to absorb its kinetic energy, most neutron shielding is is boron or a hydride – 200-300mm is enough.
The NTTR version of this rocket has on-board shielding sufficiently thick to limit payloads and ground observers to 0.05 rad per launch mission (about 20% of an annual radiation dose if you don’t fly commercial airlines, if you do then less than 1%).
Mmmm kay.
Out of curiosity how much mass can be delivered to the lunar surface? (Single stage to lunar surface)
Other question: Did you watch Destination Moon yet?
The lox-augmented nozzle is a good alternative if you are running fuel-rich in the primary combustor. However, the Isp is much lower for oxized post-combustor injection.
There are several NTP reactor architectures that work, most promising as far as availability is a high-assay LEU (19.8% enrichment) currently being worked on by NASA.
Question, Mr Bucknell: what sort of enrichment percentage would such an engine use, and would it be possible to run it on a different fissile element (say U-233)?
6.5m core is optimal until demand rises.
Thanks, you haven’t considered a slightly bigger version like 9m. I’m guessing this is the sweet spot.
Also, the costs of complex structures for aerospace have ALREADY reached less than $2000/kg for vehicle level costs (example: FH). The architecture presented is so much smaller and lighter per unit payload that the $25M is a conservative estimate.
The arguments for methane don’t hold in this architecture. The tankage is very small, so the aerodynamic and mass advantages of smaller tanks don’t offset the loss in Isp and therefore delta V. The bigger challenge is that methane doesn’t have the wide flammability limits and therefore the velocity range of airbreathing will drop by about half – reducing the Isp benefit for terrestrial launch. Lastly, any ISRU that generates methane generates hydrolox as an intermediate step.
How would the numbers work for a version the same diameter as the Starship running on methane?
I would contend that reuse with multiple vehicles is harder due to recovery performance penalties as well as reintegration challenges. The variable geometry allows high performance over a very wide Mach number range – jettisoning such a highly integrated system would likely result in a net performance penalty.
One note I might add is that half of all launch failures in the last twenty years have been due to upper stage issues – either separation went wrong or upper stage didn’t light. That is significant risk.
https://www.quora.com/What-are-the-major-causes-of-rocket-failures
In-space is right around 900s for the proposed and historical reactors.
On the other hand, this architecture pays for the reactor to be hauled up as dry mass rather than payload on another flight – so large savings right there.
OK, that plane should have been approved for production on looks alone.
Looking at the graphs, that’s an in-space Isp of about 900 right? Which would make a serious difference once you’ve got a space ship in space with a given fuel load. But you’ve still got the extra reactor mass to deal with.
I would expect it really helps if you have lots of delta V to get where you are going after LEO. The moon for example. Or Mars.
But we have John himself here now so we can get the real story.
Fundamentally the wireless power transfer physics requires large aperatures on the transmitter side and the receiver side with radio frequencies – lasers are problematic due to energy density. It is relatively unrelated to the power absorber.
Does SPSS even have to have large surface area. I wonder if coils alone can be used to generate/store power.
The challenge with that approach, as tempting as it is, is that turbojets just aren’t that powerful and turbofans are only good <M0.9. At best, you might be able to achieve 900 m/s of delta-v (say M2+gravity loss). You don’t gain much, and now you have to carry the dead weight of those turbojets with you. Definitely a deal-breaker for SSTO, though might be an acceptable cost for TSTO.
The alternative is to simply push your rocket engines hard to M1, where your air augmentation or ramjet really kicks in. Here your only cost is oversized rocket engines (which you already have) and extra fuel.
Aerojet (I think) has patents for low-pressure combustion +within+ the rocket bell as an alternative to aerospikes. The gis is that you use a high-efficiency, high-pressure vacuum optimized engine, and then at low altitudes dump low-pressure propellant into the bell and use the bell itself as a booster.
It would have been helpful if Brian had mentioned the two variants. 🙂
No idea what cost analysis was used to reach $25 million. Seems pretty optimistic to me, since this is effectively a very large jet engine, and those aren’t cheap.
Fuel is cheap, so dropping the complexity and weight of the turbofan should improve construction costs.
TSTO probably increases +total+ costs, since now you have two rockets to build. However, as we have seen with F9, it makes reusability easier. It also reduces risk, since now you can use off-the-shelf components for the second stage. It should also increase payload capacity, though that might be balanced by the increased dry mass of the interstage and additional hardware.
There aren’t any reactors available yet anyhow, but the in-space performance enhancement offers huge gains for NEO missions.
Presentation materials are online, still waiting for audio files to merge.
https://youtu.be/8pfKWDMc73o
In Atmosphere, anything with a nuclear thermal system is a no-go. Only hope for nuclear thermal based propulsion is for inter-planetary or whatever, once in space, so the magnetosphere shields us.
The Dyson-Harrop satellite scheme isn’t easily deployed currently. It may well be possible in the future.
What with all the migrations, the former articles aren’t obvious:
https://www.nextbigfuture.com/2015/07/interview-with-john-bucknell-about-his.html
https://www.nextbigfuture.com/2017/07/air-enhanced-nuclear-thermal-rocket-by-x-spacex-engineer.html
https://www.nextbigfuture.com/2017/07/john-becknell-designer-of-the-nuclear-thermal-turbo-rocket-is-answering-questions.html
https://www.nextbigfuture.com/2018/07/146443.html
EDIT: I’m an idiot, scanned right over the last line. I look forward to seeing the presentation.
I know that this isn’t really your focus, but since you’re looking into Space-based Solar Power systems I suppose I’ll ask: have you looked into Dyson-Harrop satellites at all? I know that the concept isn’t very well developed, but if it works then putting up a couple of them and some laser relay satellites to beam power back to Earth from outside the magnetosphere seems a lot less costly than putting up the giant arrays you need to harvest sunlight from closer orbits.
To address the questions, there are chemical and nuclear variants with primary core/combustor running fuel rich. The chemical variant is substantially cheaper, and enables low cost access to space.
Hot hydrogen fuel for the air-breathing is an enabler for the entire architecture, extending air breathing over 72% of the delta vee to achieve orbit.
The presentation and commentary will be posted in the next few days.
John
The idea has been around, and even used on military propulsion. The question is whether the extra thrust by pushing added air is worth the added structure and complexity. There is also a non-turbo version known as a ducted rocket or ejector rocket, where extra air flow is dragged along by the rocket exhaust. The alternative is to simply use an air-breathing first stage, with a turbojet/ramjet, and stage off a conventional rocket when you run out of air.
https://en.wikipedia.org/wiki/Air_turborocket
Would any hypothetical changes to accommodate your perceived issues in the design maintain or improve the estimated construction cost of $25m?
As is too often the case, it looks like Brian has smushed together (a technical term from literary criticism theory) two different descriptions of two different ideas.
If you google (or duckduck) “John Bucknell Turbo Rocket” you can find a video of his presentation.
The money shot is
So both a nuke and pussy version of the same thing. The nuke version scores Isp = 1695, non-nuke version Isp = 1430.
I would be very surprised if the advantage of the nuke model even paid for the extra expense, let alone the political firestorm that would result.
I thought the big difference with this one might be that it was not nuclear, but I see the “hot-hydrogen propellant” and it sounds like nuclear.
It will never get anywhere.
Could we use the same turbo-rocket concept with ‘regular’ propellants and see a similar increase in ISP?
Oooh, I like the idea of axisymmetric ram/scram-jets.
I think the rocket-fan adds additional complexity and mass that is only used for 300-600m/s.
The mass of those variable geometry inlet and throat is completely unnecessary once past the atmosphere. Does it make sense to build it as a TSTO, get rid of the additional complexity of thin-film cooling, and do stage separation at Mach 15?
No offense but this is one of those things that just isn’t going to happen. Look people will NOT be ok with a nuclear rocket that operates in the atmosphere. One screw up and a chunk of spaceship becomes a radioactive chunk hurtling into a city or something? Nope.
Out of atmosphere Sure. In it not for a while.