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The large ship that Robert Dyke is leading the design of, requires a propulsion system, capable of achieving the following - Outbound:
1) Transit from LEO to Earth Escape (3km/s);
2) dV associated with interplanetary transfer orbit between Earth and Mars (3km/s?);
3) Disposing of excess velocity resulting from entering the Martian gravity well (5.2km/s?).
Inbound would be the opposite steps 4 & 5, but entering Earth orbit will result in excess velocity of over 11km/s. Presumably, at both Earth and Mars, there would be either an initial pass through the atmosphere, losing enough energy to put the craft in a highly elliptical orbit that can gradually be circularised by successive passes through the atmosphere. This leaves steps 1, 2, 4 & 5, which must be accomplished using propulsion. This is a total dV of around 12km/s for a round trip. If chemical propellant were used, the mass ratio of the ship would be unfavourable, requiring a greater mass of propellant than the empty mass of the ship itself. This would make operating the ship costly.
The resistojet propulsion concept uses electrical heat to heat a propellant to high temperature and exhaust it through a nozzle. This is an extremely simple concept. The maximum exhaust velocity that is attainable is a function of gas temperature before expansion:
https://www.fxsolver.com/solve/
My proposed concept is for a rocket consisting of tungsten tubes set within a sintered thorium oxide block. Around the thorium, we have a wound copper coil, carrying high frequency AC. The rotating magnetic field heats the tungsten tubes to 3300°C (3600K), which is just beneath the melting temperature of the thorium oxide. We pass hydrogen gas through the tubes. Pressure before the ThO2 exit nozzle would be about 1KPa. According to the calculator, this gives an exhaust velocity of 11,880m/s, or an ISP 1211s. This is enough to fly from low Mars orbit and back again, with a mass ratio of about 2.
Additional: This thruster is a thermodynamic device. We are converting electrical energy into the heat of a gas, which then expands through a nozzle, coverting internal translational energy into kinetic energy. Overall efficiency electrical to KE is about 30%.
Last edited by Calliban (2024-01-18 09:47:28)
"Plan and prepare for every possibility, and you will never act. It is nobler to have courage as we stumble into half the things we fear than to analyse every possible obstacle and begin nothing. Great things are achieved by embracing great dangers."
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For Calliban re interesting new topic...
Best wishes for success with this topic, and for encouraging members to contribute to it's development!
Something that I am hoping will come out of the discussion is an estimate for thrust that might be produced using this system.
As a reminder, for planning, you may use 5000 metric tons as the mass of Large Ship ready to leave orbit at Earth.
(th)
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I've seen proposals for resistojets back in the 1950's and 1960's. There have been some bench test demonstrations in the 1960's. You can create a propulsive jet that way, but the electric power required to do it was always enormous! Going to high chamber pressure makes the electric arc heating way more difficult yet, but is required to make the propulsive jet efficient. That's about all that I know about it.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Aerojet Rocketdyne currently manufactures the AR-510 for satellites. It's an arcjet using high purity hydrazine monopropellant. A single thruster mass 1.6kg, power 2000 watt. A module of 4 thrusters with common power module, dry mass 17.4kg, 70 volt, average power consumption 2,200 watt, specific impulse 585 to 615 seconds, thrust 258 to 222 mN.
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You realize I said aerocapture to slow when entering planetary orbit on either end. And we now have a new way of slowing into Earth orbit: what is the ∆V from plowing through StarLink satellites?
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Hydrazine is a monopropellant that just needs a small "kick" to exothermically decompose. Milli-N of thrust from kg of equipment (NOT counting the electric power supply!!!), and requiring KW to run, does not sound all that attractive to me, except for the same sorts of applications where the other electric propulsion notions are also feasible.
If you cut the Isp by about 3 or 4 and go to higher pressure, you can get kilo-N of thrust out of those same kg of equipment, and not draw much, if any, electric power. Worst case: a heated catalyst pebble bed to decompose the hydrazine. "Kg of equipment" does not include the propellant tanks, nor did the resisto-jet number include it!
Add NTO and you get the same kilo-N of thrust at no electrical draw out of only "kg of equipment", and at about half the Isp of the resisto-jet. Hypergolic ignition.
GW
PS - I suspect the drag of plowing through Starlink satellites might be rather erratically large. (He says with a chuckle.)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Another form of ion engines when the fuel is heated and not burned to create thrust.
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