guest blogger john hare
Once a vehicle is in orbit, a high thrust to weight engine becomes a convenience rather than a necessity. For orbital transfer, Isp becomes far more important due to the cost of delivering propellant to orbit. Nuclear thermal seems like a good compromise between thrust to weight and Isp except for the political problems with getting it developed. All of the electrical propulsion system suffer from excessive weight when the power generation requirements are considered. Solar sails will be better for deep space work if they are ever developed and get appropriate markets.
There is at least one credible plan out there to deliver propulsion modules to orbit for $100.00 per pound of propellant. If this happens, then the propulsion for even a mars mission becomes economical enough that all the numbers change. If the modular propulsion modules become available at that price, then 5,000 tons of propellant with tanks and engines would cost one billion dollars delivered to orbit. That would put over a thousand tons of people, consumables, and gear in mars orbit for a fraction of one year of the NASA manned spaceflight budget.
Until some of the above items start delivery, a high Isp engine with relatively high thrust to weight compared to ion and other electric thrusters will be very attractive. I propose a 750+ second orbital transfer engine with an estimated T/W of 0.1. Or one meter per second acceleration of the tetherocket system mass with no propellant or payload. It is a mix of tether and rocket technology with some nuclear or beamed energy in the mix. I believe it could be done for an investment that would be economical compared to most of the alternates I am aware of. I believe the performance compromises would be favorable compared to most proposed systems.
Current tether technology can easily get 3,000 meter per second tip speeds on a rotating tether in vacuum. Current hydrogen oxygen rockets can easily get 4,500 meter per second exhaust velocities in vacuum. If an axle anchors the center of the tether to the vehicle with rockets at the spinning tips, then the rockets can pulse during the 10% of the rotation when the vectors line up. When the rockets pulse, the 4,500 meters per second exhaust velocity is added to the 3,000 meters per second tether velocity to give a total exhaust velocity of 7,500 meters per second relative to the vehicle.
The catch (you know there is one) is that the tether must be powered to reach and keep the 3,000 meter per second tip velocity. It also must transfer the momentum of the rocket pulses to the ship to turn them into propulsion. This will take a considerable amount of power. Nuclear engines would be a good choice with nuclear steam driving the tether. The hydrogen and oxygen could provide the cooling cycle for the nuclear engine before injection into the pulsed thrusters. The nuclear engine would be regeneratively cooled this way with the waste heat providing propulsion gain.
If this propulsion system can be built, it would be somewhat useful for geosats and lunar missions. It would be very useful for NEO, mars, and farther missions. For a lunar mission, it would cut the mass ratio from ~2.4 down to ~1.7 starting in LEO and ending in LLO. For asteroid and mars missions, it could cut travel time by more than half, or increase the available mass to the destination.
In the sketch above, blue represents the tether being spun counterclockwise with thrusters only burning from the 10 o’clock to the 8 o’clock position. Yellow is the clockwise tether with a burn from the 2 o’clock to the 4 o’clock positions. With the start and stop transients, true thrust is about a 10% duty cycle. Direction of ship travel is up the page.
I’ve read that current 3,000 m/s tethers can handle tip loads of 1% of their own mass without problems. Each tether has two thrusters firing a total of 20% of the time. A 100 lb tether can hold two 1 lb mass thrusters. A 1 lb mass thruster can have a thrust to weight of 100. Each 102 lbs of tether and thruster will average 20 lbs of thrust. The nuclear or beamed energy engine that drives the tethers should mass about the same. A bare engine would accelerate at about 1 m/s^2. If the engine mass is 10% of the total ship mass, then acceleration is about 100 cm/s^2. It would take half a day to reach escape velocity from LEO. There would be some gravity losses taking this much time to leave orbit. This thing would be more useful in longer range operations.
Feeding and igniting the thrusters seems more difficult than it actually is. If a fine propellant mixture is sprayed in front of each thruster as it approaches the pulse point, the propellants will impact the thruster chamber at 3,000 m/s. The impact will heat the mixed propellants well past the temperature of ignition. The heat generated by the impact boosts the thruster Isp by increasing the mix start temperature. The reaction is that of a pulse detonation engine. Ignoring the performance claims of pulse detonation engine advocates, the hydrogen/oxygen thruster can reach 4,500 m/s exhaust velocities plus some gain from the very high pressure and high expansion ratio possible here. If this reaction can be brought to service, I believe it quite possible that net Isp will be in the 850 range with system thrust to weight double what I am suggesting here.
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