NASA has small $18.8 million nuclear thermal rocket research project

Last year, NASA partnered with BWXT Nuclear Energy Inc. for an $18.8 million contract to design a reactor and develop fuel for use in a nuclear-thermal propulsion engine for deep-space travel.

“Significant advances in material research and technology development have allowed for new materials to be considered for the critical components of the reactor,” said BWXT’s Cirtain.

Given its experience in developing and delivering nuclear fuels for the U.S. Navy, BWXT will aid in the design and testing of a promising, low-enriched uranium-based nuclear thermal engine concept and “Cermet” — ceramic metallic — fuel element technology. During this three-year, $18.8-million contract, the company will manufacture and test prototype fuel elements and also help NASA properly address and resolve nuclear licensing and regulatory requirements. BWXT will aid NASA in refining the feasibility and affordability of developing a nuclear thermal propulsion engine, delivering the technical and programmatic data needed to determine how to implement this promising technology in years to come.

The overall goal of this game changing technology project is to determine the feasibility and affordability of an LEU based NTP engine with solid cost and schedule confidence. Initial project goals are to demonstrate the ability to purify tungsten to a minimum of 90 percent purity and determine the production costs at that purity level; to determine the technical and programmatic feasibility (pre-phase A level) of an NTP engine in the thrust range of interest for a human Mars mission; and to determine the program cost of an LEU NTP system and the confidence level of each major cost element.

Because ground testing an NTP engine is more difficult than chemical engines, the project will examine the feasibility of contained engine testing within environmental and safety guidelines. Initial testing of fuel element materials will include non-nuclear tests in near prototypic conditions. An advantage of an LEU-based system is the possibility of total containment testing at a conventional propulsion test facility such as Stennis Space Center, which further reduces cost and complexity.

With the potential to provide high thrust at over twice the specific impulse of the best chemical engines, there is no doubt that the capability provided by NTP is a game changer for space exploration. The Game Changing Development (GCD) Program investigates ideas and approaches that could solve significant technological problems and revolutionize future space endeavors.

A large nuclear thermal rocket could outperform chemical rockets.

With refueling in high orbits a SpaceX BFR can get a delta-v (change in velocity up to 9 kilometers per second).

Space Missions to Mars up to now have been small satellites where the entire mission was launched from Earth. The delta-V has been about 3.2 to 3.5 kilometers per second.

By refueling the SpaceX BFR in orbit and assembling a few stages in high orbit, a large chemically powered space mission can get up around 9.0 kilometer per second delta-V.

Higher delta-V opens up faster trajectories and faster missions to Mars and other planets. SpaceX BFR will be raising the bar for nuclear thermal rockets. Nuclear thermal rockets will need to do more than just beat the 3.5 kilometer per second of small upper stage chemical rockets.

A Large reusable Air enhanced nuclear thermal rocket design like the one designed by John Bucknell will be needed to surpass the SpaceX BFR

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. ISP is a rocket fuel efficiency measure like miles per gallon on the ground.

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.

Comment question
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.

Answer Bucknell.

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).

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