NASA Funds Breakthrough Nuclear Rocket

A breakthrough pulsed plasma rocket is getting NASA phase 2 innovative concepts funding. It will have 15 times the efficiency of chemical rockets and with high levels of thrust at the level of a medium sized chemical rocket.

Howe Industries is currently developing a propulsion system that may generate up to 100,000 Newtons of thrust with a specific impulse (Isp) of 5,000 seconds. Chemical rockets have about 300-400 specific impulse which means less fuel efficiency so that the time an engine can be used before running out of fuel is less.

The SpaceX Merlin chemical rocket engine used on a Falcon 9 has about 340,000 newtons of thrust. A pulsed plasma fission fusion engine will enable spaceships to go ten to thirty times faster than chemical rockets.

The Pulsed Plasma Rocket (PPR) is originally derived from the Pulsed Fission Fusion concept, but is smaller, simpler, and more affordable. The exceptional performance of the PPR, combining high Isp and high thrust, holds the potential to revolutionize space exploration. The system’s high efficiency allows for manned missions to Mars to be completed within a mere two months. Alternatively, the PPR enables the transport of much heavier spacecraft that are equipped with shielding against Galactic Cosmic Rays, thereby reducing crew exposure to negligible levels. The system can also be used for other far range missions, such as those to the Asteroid Belt or even to the 550 AU location, where the Sun’s gravitational lens focuses can be considered. The PPR enables a whole new era in space exploration.

The NIAC Phase II conceptual studies will receive up to $600,000 to continue working over the next two years to address key remaining technical and budget hurdles and pave their development path forward. When Phase II is complete, these studies could advance to the final NIAC phase, earning additional funding and development consideration toward becoming a future aerospace mission.

The NIAC Phase I study focused on a large, heavily shielded ship to transport humans and cargo to Mars for the development of a Martian base. The main topics included: assessing the neutronics of the system, designing the spacecraft, power system, and necessary subsystems, analyzing the magnetic nozzle capabilities, and determining trajectories and benefits of the PPR. Phase II will build upon these assessments and further the PPR concept.

In Phase II, they plan to:

* Optimize the engine design for reduced mass and higher Isp
* Perform proof-of-concept experiments of major components
* Complete a ship design for shielded human missions to Mars

11 thoughts on “NASA Funds Breakthrough Nuclear Rocket”

  1. My take is there is an operating fission reactor powering the Brayton engine, and providing a high neutron flux cylindrical volume. I’d guess it uses heavy water as a moderator.

    When thrust is desired pellets of something fissionable maybe highly enriched uranium hydride pellets at high speed are introduced into the center of the reactor. The pellet explodes into plasma, and the plasma is directed out the rear by magnetic fields. I think I would incorporate a metal shield that would drop in line after the pellet passed.

  2. Make-a-wishonium … sounds like.

    Here’s one to remember: the energy contained in the exhaust plume — kinetic energy — has a very straight-forward physics equation

    [1.1]  Ek = ½mv²

    The [v] part is related to [Isp in seconds] by

    [2.1]  v = G₀ • Isp
    [2.2]  v = 9.81 • Isp

    No biggie … just substitute [2.1] in to [1.1]

    [1.2]  Ek = ½mG₀²Isp²

    Yah, yah, math. But it is SIMPLE math: we’ve got a [5000 sec] Isp quotation. Using 1 kilogram of effluent

    [1.3]  Ek = ½(1 kg)(9.81 × 5000)²
    [1.4]  Ek = 1,203,000,000 joules/kg
    [1.4]  Ek = 335 kWh/kg

    So, at the very minimum, it takes MORE than 335 kWh or 1.2×10⁹ joules of investment energy into EVERY kilogram of effluent, to achieve the 5,000 Isp thrust rating.

    Last I heard, that is a LOT of energy. Now one can tame this wild animal by metering out only a modest amount of kilograms per second, or per day. Then the power demand is more approachable. Let’s say “1 kg/day”

    [3.1]  Δm/Δt = 1 kg / (1 day × 24 hr × 60 min × 60 sec)
    [3.2]  Δm/Δt = 11.6 milligrams per second

    [4.1]  power = 11.6×10⁻⁶ kg × 1.2×10⁹ J/kg (watts)
    [4.2]  power = 14,000 W

    Now that’s the joule-seconds (watts) in the exhaust plume in kinetic energy. Real thrusters might be characterized as having a 80% or so (really good) efficiency, so

    [5.1]  input power = power / efficiency
    [5.2]  input power = 14,000 W ÷ 0.80
    [5.3]  input power = 17,500 W

    See? A perfectly reasonable amount of power at 1 kg/day of reaction mass burn rate.

    But, if your rocketeer sensibilities are still reading and working, you’d ask, but what does this 11.6 milligrams per second or 1 kg per day give us in actual thrust? Well, that’s still not hard:

    [6.1]  impulse = momentum = mv = Δm/Δt • t • (G₀ • Isp)
    [6.2]  impulse = 11.6×10⁻⁶ kg/s × 1 sec × 9.81 × 5000
    [6.3]  impulse = 0.57 newtons

    Well, that’s pathetic! Clearly it needs to be a WAY bigger number for a real-world space ship having a few space-farers aboard, along with a year’s worth of vittles, oxygen, water, change-of-undies, washing machines and detergents. Tons of stuff per astro-nut. Tons. And of course a protective space-ship to hold it, and keep most of the bad space-radiation from killing everyone before arrival. Tons and tons and tons and tons.

    Let’s say 25 tons. Nice round number. We need (say, this being napkin math) 100 km/s of ΔV in (say) 30 days. 3 m/s per day. 3000 m/s. Using

    [7.1]  ΔV/Δt = p / m₀ where [p = the Ek/t of the exhaust]
    [7.2]  ΔV/Δt = p ÷ 25,000 kg … move stuff about
    [7.3]  3000/(24 × 60 × 60) × 25,000 = p
    [7.4]  p = 868 newtons

    Well, there we are. My intuition was spot-on.
    Need a LOT more than 0.57 newtons.
    In fact it is multiplicative …

    [8.1]  m / day = 1 kg × (868 ÷ 0.57)
    [8.2]  m rate = 1,570 kg/day

    And the specific energy is likewise 1,570 times 17,500 watts or 27,000,000 watts.

    27 megawatts of energy input.
    That has to be one heck of an onboard power generator.
    Nuclear, of course.
    And at an expected efficiency of maybe only 35%, well over 100,000,000 watts thermal. Nice!

    Anyway, just a little napkin astrophysics.

    To [Elton John] … “Rocketman, da-de-da-da-daaaa-da”

    ⋅-⋅-⋅ Just saying, ⋅-⋅-⋅
    ⋅-=≡ GoatGuy-who-does-rocket-calculations ✓ ≡=-⋅

  3. Musk and his cow gas rockets are not the way forward to travel to and colonize Mars. We are going to need nuclear to travel to and from Mars quickly. To travel 6 months to reach Mars by Starship is stupid slow. To keep LOX and LNG liquid during space travel will be a major “head-ach”. If we have nuke ships that can go 250K+ MPH than we need to exploit the speed to mine asteroids for metals. That’s where the profits lay, not Musk’s “Total Recall” wet dream.

    • The SpaceX methane-oxygen rocket system is pretty good, for the gravitationally hard part of the flight from Terra to Mars … namely, LIFT OFF. Lest we forget, NASA used specially refined kerosene and oxygen to get the Atlas V off Planet Dirt, with EVERYTHING aboard for the trip to Luna … and back! Kerosene. Lamp oil. Heavier than Diesel fuel.

      At the other extreme, NASA went on to use hydrogen-oxygen, to minimize the molecular weight of the effluent gas, and derive ostensibly highest ISP for the kilograms of fuel lofted. Only problem was, liquid hydrogen is the opposite of compact. 70 kg per cubic meter. Big, really big tank. External tank. Methane is really a reasonable compromise between carbon-rich kerosene (and all its rocket-engine complications) and stupid-fluffy hydrogen. 650 kg/m³ for liquid methane. 25% less than Diesel or kerosene.

      Once the stupendous-burn-rate off-Earth lifting is done, per the genius of SpaceX, just turn around the boosters, and land them back on the ground with less than 6% of the original fuel in reserve for the operation. Keeps costs down.

      Build up with multiple missions all the crâhp up there that’d be needed to fuel-up, provision and refurbish the inter-planetary orbiters that’d take us to Mars and the Asteroids and other delicious destinations in the least amount of reasonable time.

      ⋅-⋅-⋅ Just saying, ⋅-⋅-⋅
      ⋅-=≡ GoatGuy ✓ ≡=-⋅

  4. Ooooooh, “Fourier series approximation control drums with mini sub cores.” Like a doodle sketch made in a boring class. I have to remember/remind myself that everything is new to the Next Generation.

  5. If Mars is at closest 40 million miles from Earth and this thing reaches a max stated velocity of 500k mi/hr, how long would it take, with acceleration and deceleration of a spacecraft for a Hohman Transfer?

    • A Hohman Transfer, by definition, is a minimum delta vee transfer orbit. For Earth to Mars, the time would be roughly nine months. That’s definitely NOT the type of orbit a nuclear rocket discussed her would take.

      For a rocket with a maximum delta vee of 500,000 mph, a first order approximation for the Earth – Mars tranfer would be maximum acceleration to 250,000 mph, coasting for most of the distance at that speed, and then maximum deceleration to Mars orbit. To cover 40 million miles, that’s a maximum time at null-gee cruise of 160 hours, or a little under a week. But depending on the thrust to mass ratio of the rocket — which determines its maximum acceleration, it would probably need a couple of days accelerating to cruise speed, and another couple of days decelerating to Mars orbit insertion. So call ita little over a week for the overall trip time.

      • Could they use a laser beam ‘Repercussion’ effect, lasers can PUSH, to contain the exhaust of this Pulsed Plasma Rocket, so that its propulsion is contained in a very long, straight-line action/reaction alignment tapered cone of projected laser light, instead of much of its thrust being displaced sideways by the vacuum of outer space?

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