John Becknell designer of the Nuclear Thermal Turbo Rocket (NTTR) and x-senior engineer on the Space Raptor engine is answering questions at nextbigfuture. He is answering on the previous article on his air enhanced nuclear thermal rocket that has 5 times the ISP of a chemical rocket and 10 times the payload. The increased payload is from having ten times higher payload fraction.
In 2015, Bucknell presented the Nuclear Thermal Turbo rocket which added air-breathing to a nuclear thermal rocket. Bucknell design would have 1664 ISP. 60% more than the best prior nuclear thermal rocket designs.
At next months’s Icarus Interstellar’s Starship Congress in Monterey, CA, John Bucknell will present a mission analysis using the NTTR architecture titled “Single Stage to Orbital Habitat”.
Question set 1. Is the rocket reusable? Is it intended for reuse?
What happens if the rocket explodes in flight? (how strong is fuel containment)
How much gamma/neutron radiation near the launch pad at takeoff?
Any numbers for a single stage to Mars surface mission profile?
How long to get 20,000 kg to a gravitational lens location?
Any good links on MITEE? Web searches last week didn’t turn up much.
Yes, the NTTR architecture is intended to lower launch costs by both reducing cost/payload AND full reuse.
Exploding is not a thing with a monopropellant rocket (how often do aircraft explode?). As for fuel containment/reactor structure – if the rocket were to have a RUD event the core/shielding is largely metallic with a tungsten carbide gamma shield (six of the seven tons of core mass). It would put a large dent in anything it landed on, but that is a risk with any rocket (ie why launches are over water).
The radiation shielding included in the rocket is probably overkill – it keeps exposure to both payloads and launch site below terrestrial background radiation (0.2 rad/y) if less than five flights a year are flown. Gamma shielding is fully surrounding the reactor (tungsten carbide), and lithium hydride for neutron fore for payload protection. Details in the paper.
Single stage to Mars surface aren’t quite there yet, but to Mars orbit are analyzed in the 2015 SC pres. Larger rockets help this, ~15% payload fraction is predicted for a scaled up 5,000MW variant (same as Lunar surface). Have to work on gravitational lens mission.
I had to go to Internet Archive to get MITEE data. Plus Ultra Technologies closed up shop a few years ago. I’ll see if I can find the link.
Question 2. How is exhaust radioactivity prevented or dealt with? Would I be correct to say that the hydrogen propellant doesn’t stay long enough in the reactor to be transmuted to tritium, and any fission fragments are contained by the fuel cladding?
Answer. What exhaust radioactivity? There are no fission fragments with this fuel arrangement as core temperature is kept low for effectively zero material loss. Hydrogen spends milliseconds in the core, so mass of tritium generated is negligible.
3. What are the risks for catastrophic failures inside the atmosphere, and how would they be dealt with?
See answer 1 above
4. Would using water instead of hydrogen as propellant make sense? As I understand, the heavier oxygen should result in lower Isp, but on the other hand, water is significantly denser than liquid hydrogen, so more can be carried in the same volume.
Water is awful as a terrestrial launch propellant, only 200s ISP. And it isn’t reactive with oxygen so no chemical boost to propellant temp. ISP is king for launch payload fraction – anything less than 800s doesn’t even make orbit (without staging) since the reactor mass is so high.
5. I assume the rocket is intended to be reusable. Is it intended to RTLS and land on its tail like SpaceX rockets? The illustrations include what looks like small landing legs.
My [Bucknell] concept model is to introduce the rocket propulsion cycle, so yes the landing legs are underdeveloped.
As for the nozzle, sliding cone/cylinder structures using the centerbody are pretty well known solutions. But this NTTR proposal seems to use some sort of widemouth inlet that expands beyond initial diameter, which seems difficult compared to say an expandable/extensible nozzle. Sliding cone inlet seems easier.
The sliding interface doesn’t give variable shock angle capability which is necessary to provide the early ‘start’ and wide free stream velocity range (Mach 3-14). Think of it is a rear-hinged ramp, arranged in a circular pattern. Most inlets are ‘2D’ to achieve a variable inlet with a forward hinge.
Other Bucknell comments.
Hydrogen pumped up to plasma would help in-space ISP. NTER mentioned below is an attempt to help that – but vehicle mass goes through the roof.
The variable geometry inlet is actually a straight inversion of the variable geometry nozzles used in jet aircraft (look up Pitch/Yaw Balance Beam Nozzle), with one degree of freedom removed (throat area variable).
The Turbo-Inductor from the NTER is a great idea, but generally only helps the in-space thrust by raising the propellant temp (combustion is already temp-limited). If you look at the NTER implementation, it is a huge heat exchanger – again far too heavy for a launch vehicle.
Ground testing of MITEE is actually easier than most – since each individual pressure tube fuel elements is a full-scale rocket (the 1,000 MW variant is about 169 individual fuel elements).