The Nuclear Thermal Turbo Rocket - A Conceptual High-Performance Earth-to-Orbit Propulsion System by John Bucknell
A new propulsion concept called the Nuclear Thermal Turbo Rocket (NTTR) is proposed for Earth to Orbit applications. The NTTR utilizes a nuclear fission reactor to thermally heat hydrogen propellant into a rocket plenum. The rocket nozzles are located at the tips of a variable pitch thrust fan connected to the plenum by passages in the fan blades, and each nozzle is a linear aerospike on the trailing edge of the blade. The thrust fan is located in a duct such that the heated hydrogen propellant is combusted with ambient sourced oxygen to augment the rocket thrust. The fan is of variable pitch to provide maximum thrust for varying inlet velocity. The duct has a variable geometry inlet, able to provide appropriate mass flow and compression to the combustor throughout the trajectory, and a variable geometry outlet to provide appropriate nozzle area for maximum thrust. The rocket nozzles act as propellant injectors during the airbreathing portion and pure rockets during low atmospheric density portions, with the NTTR utilizing a single gas path from launch to orbital velocity. The propulsion concept is of high performance and is able to transport more than 50% mass fraction in a Single Stage to Orbit (SSTO) via an air-breathing rocket trajectory with intended complete reusability. Payload fractions of up to 19% are predicted (inert mass includes reactor radiation shielding) due to a mission average Specific Impulse (Isp) of 1,662 seconds
Here is information from an email interview with John Bucknell.
Question 1. Can you list out your modifications compared to older designs and experiments ?
Background to question 1 -
The NERVA experiments had an ISP of about 875, and the thinking was they could have been upgraded to 975
The Timberwind design (1987-91) in theory could reach 1000 ISP
And some current designs are at 925 ISP.
There were nuclear light bulb - gas core design open cycle design with ISP in the 3000 to 5000 range and closed cycle 1500-2000 ISP.
Answer 1 from John Bucknell
The pure rocket portion of the system is pretty conventional - I used an off-the-shelf design reactor (MITEE) with peak propellant temps limited to about 2200 deg C so as to limit fuel element erosion that starts at about 2500 deg C. Mass loss of the fuel element at 2750 deg C was a fraction of a percent per 100 hours of operation (ie not much). The Isp as a pure rocket is 890 seconds in vacuum. When in airbreathing mode the exit temp drops since so much more fuel is being pushed through the core - with around 1000 deg C propellant temps. So no exotic designs needed, it's the airbreathing that ups the Isp.
Question 2. Also, if Spacex gets reusable stages then how does yours reusable nuclear thermal rocket compare ?
Answer 2 from John Bucknell
As compared to SpaceX's reusability - same motivations. However, a SSTO can land, refuel and launch again whereas the F9R needs to be reintegrated. And the upper stage doesn't have the mass budget (yet) for propulsive recovery. But the big kicker is payload fractions - my design is only 15% of the GLOW of a F9R for the same payload (at the low end of estimates - top end is 1.5x payload), so the rocket is far simpler and easier to build. And bigger rockets generally have better payload fractions - so a scaled up version could get Saturn V sized payloads at still only 60% of the GLOW of a F9R.
Question 3. Also, no one has gotten scramjets working. How difficult will it be to get some of the pieces you had talked about working and then integrated ? Is it somehow easier to develop and integrate what you are doing ?
Answer 3 from John Bucknell
Actually, yes - scramjets have been demonstrated experimentally in the same flight regime as I'm proposing (albeit performance data not publicly disclosed). And that is with hydrocarbon fuel which doesn't burn well. The superheated hydrogen makes combustion far simpler, which is where all the other designs struggle. No doubt much work to be done in this area - but extensively studied (the demo vehicle technologies are over twenty years old).
Question 4. People talk about the safety concerns. What if your system blew up ? How would the shielding survive the explosion and the crash into the ground
Answer 4 from John Bucknell
Yes, safety. Firstly, the greatest danger is dropping a 16klb tungsten carbide radiation shield on your head (same with all rockets, the engines of a F9R weigh a collective 18klb of high strength metal). Since there is no oxidizer but air, there really isn't an explosion risk like other rockets more like a loss of propellant accident (ie no thrust). The radiation concerns assume you can't shut the reactor down - a risk with all reactors, which is why we talk about defense in depth and redundancy in the control rods/neutron poison actuation such that the reactor does power down in case of an emergency. The radiation shield is overkill even while the reactor is operating - payloads receive less than terrestrial background radiation.
The development time/effort is of course dependent on the cost/benefit analysis. I think the missions this rocket can do (ie single stage to the moon) can be done far more cheaply than anything else available.
The only real mission this excels at is terrestrial launch - once in orbit other nuclear rockets perform just as well. Having a rocket able to get large masses in orbit is the first step in my mind.
Here is a copy of the AIAA presentation
Nuclear thermal rockets already offer the highest Isp of launch-capable pure rocket propulsion systems, whereas launch to hypersonic turbine combined cycle systems offer far higher Isp so it is logical to attempt to integrate those cycles.
As the propellant of NTRs is superheated hydrogen fuel, introducing an oxidizer can enhance performance through combustion - raising exhaust temperature/pressure and mass flow
Combined cycle augmentation of NTRs have been proposed to enhance T/W and Isp with the LA-NTR and NEAR cycles
This design has a single gas path rocket fan, ramjet, scramjet and pure rocket combined cycle propulsion system
* Thrust augmentation from Mach 0-8 by compressing ambient air into combustor
* Fan is a variable pitch Supersonic Through Flow Fan (STFF) operated at Mach 1.5 tip speed
* Hub diameter common with core, 0.85 hub/tip ratio
* 116, 300x300mm 6% thickness semi-circular arc profile blades
* Blades duct hydrogen to two linear aerospike nozzles on each trailing edge (0.15x298mm)