High-temperature electrothermal propulsion system
In various embodiments, an electrothermal propulsion system for imparting velocity to a spacecraft features a joint-free monolithic integrated thruster component including an inlet section, a resistively heated heat-exchanger section, and a nozzle section.
1. An electrothermal propulsion system for imparting velocity to a spacecraft, the electrothermal propulsion system comprising:
a joint-free, monolithic, integrated thruster component defining:
an inlet section for receiving fluidic propellant,
a heat-exchanger section for heating the fluidic propellant received from the inlet section, thereby forming heated propellant, the heat-exchanger section comprising four or more concentric tubes fluidly and electrically connected to each other, and
a nozzle section for receiving the heated propellant from the heat-exchanger section and expelling the heated propellant, wherein the four or more concentric tubes of the heat-exchanger section form a single fluidic path for flow of the heated propellant to the nozzle section; and
a power source for applying an electrical potential between the inlet section and the nozzle section to resistively heat the heat-exchanger section to an operating temperature of 3000 K or higher,
wherein (i) no joints are present between any of the inlet section, the heat-exchanger section, and the nozzle section of the joint-free, monolithic, integrated thruster component and (ii) a wall thickness of each tube of the four or more concentric tubes is no more than 800 μm.
2. The electrothermal propulsion system of claim 1 , wherein the joint-free, monolithic, integrated thruster component comprises one or more refractory metals.
3. The electrothermal propulsion system of claim 1 , wherein the joint-free, monolithic, integrated thruster component is fabricated by an additive manufacturing technique.
4. The electrothermal propulsion system of claim 3 , wherein the additive manufacturing technique comprises layer-by-layer melting of a metal wire.
5. The electrothermal propulsion system of claim 3 , wherein the additive manufacturing technique comprises layer-by-layer melting or sintering of metal powder disposed in a powder bed.
6. The electrothermal propulsion system of claim 1 , wherein a first portion of the heat-exchanger section has a first composition and a second portion of the heat-exchanger section has a second composition different from the first composition.
7. The electrothermal propulsion system of claim 1 , wherein a composition of at least a portion of the heat-exchanger section is graded.
8. The electrothermal propulsion system of claim 1 , wherein the joint-free, monolithic, integrated thruster component defines one or more seamless bridges electrically connecting together at least two tubes of the plurality of four or more concentric tubes of the heat-exchanger section.
9. The electrothermal propulsion system of claim 1 , wherein at least two tubes of the four or more concentric tubes of the heat-exchanger section are seamlessly joined together at ends thereof via a connecting section defining one or more hollow channels therethrough for flow of the fluidic propellant from one tube of the at least two tubes to the other tube of the at least two tubes.
10. The electrothermal propulsion system of claim 1 , wherein the nozzle section has a diverging half-angle ranging from 10° to 20°.
11. The electrothermal propulsion system of claim 1 , wherein wall thicknesses of at least two tubes of the four or more concentric tubes are different from each other.
12. A method of controlling motion of a spacecraft, the method comprising:
supplying a fluidic propellant to an inlet section of a joint-free, monolithic, integrated thruster component, the joint-free, monolithic, integrated thruster component defining, without any joints therebetween, the inlet section, a nozzle section, and a heat-exchanger section fluidly coupled to the inlet section and the nozzle section;
applying an electrical potential across at least a portion of the joint-free, monolithic, integrated thruster component to resistively heat the heat-exchanger section to an operating temperature of 3000 K or higher, wherein (i) the heat-exchanger section comprises a plurality of four or more concentric tubes fluidly and electrically connected to each other, (ii) a wall thickness of each tube of the four or more concentric tubes is no more than 800 μm, and (iii) the fluidic propellant received from the inlet section is heated only by the heat-exchanger section, thereby forming heated propellant; and
expelling only the heated propellant from the nozzle section to impart velocity to the spacecraft.
13. The method of claim 12 , wherein the joint-free, monolithic, integrated thruster component comprises one or more refractory metals.
14. The method of claim 12 , further comprising, prior to supplying the fluidic propellant to the inlet section of the joint-free, monolithic, integrated thruster component, fabricating the joint-free, monolithic, integrated thruster component by an additive manufacturing technique.
15. The method of claim 14 , wherein the additive manufacturing technique comprises layer-by-layer melting of a metal wire.
16. The method of claim 14 , wherein the additive manufacturing technique comprises layer-by-layer melting or sintering of metal powder disposed in a powder bed.
17. The method of claim 12 , wherein a first portion of the heat-exchanger section has a first composition and a second portion of the heat-exchanger section has a second composition different from the first composition.
18. The method of claim 12 , wherein a composition of at least a portion of the heat-exchanger section is graded.
19. The method of claim 12 , wherein the joint-free, monolithic, integrated thruster component defines one or more seamless bridges electrically connecting together at least two tubes of the four or more concentric tubes of the heat-exchanger section.
20. The method of claim 12 , wherein at least two tubes of the four or more concentric tubes of the heat-exchanger section are seamlessly joined together at ends thereof via a connecting section defining one or more hollow channels therethrough for flow of the fluidic propellant from one tube of the at least two tubes to the other tube of the at least two tubes.
21. The method of claim 12 , wherein the fluidic propellant consists of one or more inert gases.
22. The method of claim 12 , wherein the fluidic propellant consists of xenon gas.
23. The method of claim 12 , wherein the four or more concentric tubes of the heat-exchanger section form a single fluidic path for flow of the heated propellant to the nozzle section.
24. The method of claim 12 , wherein the nozzle section has a diverging half-angle ranging from 10° to 20°.
25. The method of claim 12 , wherein wall thicknesses of at least two tubes of the four or more concentric tubes are different from each other.