Method and apparatus for small satellite propulsion

Patent metadata

Field	Value
Publication number	US-9228570-B2
Application number	US-201113578797-A
Country	US
Kind code	B2
Filing date	Feb 16, 2011
Priority date	Feb 16, 2010
Publication date	Jan 5, 2016
Grant date	Jan 5, 2016

Abstract

In a propulsion system, an electrohydrodynamic (EHD) body force is used to control the flow of a propellant through a micro channel, expansion slot, plenum chamber, or other flow region and thereby increase the specific impulse created by a propulsion system. In an embodiment, a plurality of electrodes are arranged and powered to create a plasma discharge, which can impart an EHD body force to a fluid. Various configurations of electrodes can be used to control the flow of the fluid into, out of, or through the flow region. In an embodiment, the use of EHD body forces can reduce, or substantially eliminate, shear forces on the surface of a plenum chamber, micro channel, or expansion slot of the propulsion system, resulting in a smooth flow of the propellant and increased thrust.

First claim

The invention claimed is:

1. A propulsion system for space application comprising:
a thruster comprising:
a slot shaped micro channel having an inlet and an outlet which define a thrust axis along a downstream direction, the micro channel comprising a first wall and a second wall that faces and is opposite the first wall, wherein the micro channel is configured to receive a propellant at the inlet and to exhaust propellant at the outlet;
a first electrode pair comprising a first electrode and a second electrode, the first electrode mounted on or embedded in the first wall and the second electrode mounted on or embedded in the second wall, the second electrode being downstream of the first electrode;
a power supply configured to provide a first voltage between the first electrode and the second electrode to create a first glow discharge in the propellant, the power supply and first electrode pair further configured to accelerate the propellant via a first electrohydrodynamic effect along the thrust axis to generate thrust.

2. The satellite propulsion system of claim 1 , further comprising: an expansion slot at the outlet, wherein the expansion slot is configured to expand and accelerate the propellant.

3. The satellite propulsion system of claim 1 , wherein the first wall and the second wall diverge in the downstream direction.

4. The satellite propulsion system of claim 1 , wherein the first voltage alternates at a radio frequency.

5. The satellite propulsion system of claim 1 , wherein the first electrode and the second electrode are separated from each other by an electrically insulating solid material.

6. The satellite propulsion system of claim 1 , wherein at least one of the first electrode and the second electrode is electrically insulated from the propellant.

7. The satellite propulsion system of claim 1 , wherein at least one of the first electrode and the second electrode is in electrical contact with the propellant.

8. The satellite propulsion system of claim 1 , further comprising: a plenum chamber configured to provide propellant to the inlet.

9. The satellite propulsion system of claim 8 , wherein the plenum chamber is made of TEFLON™.

10. The satellite propulsion system of claim 1 , further comprising a second electrode pair comprising a third electrode and a fourth electrode, the fourth electrode being upstream or downstream of the third electrode, wherein the third electrode is mounted on or embedded in the first wall and the fourth electrode is mounted on or embedded in the second wall, wherein the power supply is further configured to provide a second voltage between the third electrode and the fourth electrode to create a second glow discharge in the propellant, the power supply and second electrode pair are further configured to accelerate the propellant via a second electrohydrodynamic effect along the thrust axis to generate thrust.

11. The satellite propulsion system of claim 10 , wherein the first voltage and the second voltage alternate at a radio frequency.

12. A method of propelling an apparatus in space having the propulsion system for space application of claim 1 comprising the steps of:
providing propellant at the inlet;
applying the first voltage between the first electrode and the second electrode thereby creating the first glow discharge in the propellant and accelerating the propellant along the thrust axis to provide the thrust via the first electrohydrodynamic effect.

13. An apparatus comprising the propulsion system for space application according to claim 1 , for propelling the apparatus.

14. The propulsion system for space application according to claim 1 , further comprising a magnet configured to apply magnetohydrodynamic body force to the propellant.

15. The satellite propulsion system according to claim 1 , wherein the first pair of electrodes and power supply are further configured to reduce shear forces on at least one of the first wall and the second wall due to the flow of the propellant.

16. The propulsion system for space application to claim 1 , wherein the propellant is an electrically non-conducting fluid.

17. The propulsion system for space application according to claim 1 , wherein the propellant is an electrically conducting fluid.

18. The propulsion system for space application according to claim 1 , wherein the thruster is configured to produce thrust in the range of 0.8 to 1.7 mN.

19. The propulsion system for space application according to claim 1 , wherein the propellant comprises one or more of the following: Helium gas, Argon gas, Nitrogen gas, and water vapor.

20. The propulsion system for space application according to claim 1 , wherein the thruster is made of Silicon, a Dioxide compound, and a Nitride compound.

21. A propulsion system for space application comprising:
a thruster comprising:
a chip having a slot shaped channel, the channel having an inlet and an outlet which define a thrust axis along a downstream direction, the channel comprising a first wall and a second wall that faces and is opposite the first wall, wherein the channel is configured to receive a propellant at the inlet and to exhaust propellant at the outlet;
a first electrode pair comprising a first electrode and a second electrode, wherein the first electrode pair is mounted on or embedded in one of either the first wall or the second wall, the second electrode being downstream of the first electrode;
a heating element, mounted to the chip and configured to heat the chip;
a power supply configured to provide a first voltage between the first electrode and the second electrode to create a first glow discharge in the propellant, the power supply and first electrode pair further configured to accelerate the propellant via a first electrohydrodynamic effect along the thrust axis to generate thrust.

22. The propulsion system for space application of claim 21 , further comprising: an expansion slot at the outlet, wherein the expansion slot is configured to expand and accelerate the propellant.

23. The propulsion system for space application of claim 21 , wherein the first wall and the second wall diverge in the downstream direction.

24. The propulsion system for space application of claim 21 , wherein the first voltage alternates at a radio frequency.

25. The propulsion system for space application of claim 21 , wherein the first electrode and second electrode are separated from each other by an electrically insulating solid material.

26. The propulsion system for space application of claim 21 , wherein at least one of the first electrode and the second electrode is electrically insulated from the propellant.

27. The propulsion system for space application of claim 21 , wherein at least one of the first electrode and the second electrode is in electrical contact with the propellant.

28. The propulsion system for space application according to claim 21 , wherein the propellant is an electrically non-conducting fluid.

29. The propulsion system for space application according to claim 21 , wherein the propellant is an electrically conducting propellant.

30. The propulsion system for space application according to claim 21 , wherein the thruster is configured to produce a thrust in the range of 0.8 to 1.7 mN.

31. The propulsion system for space application according

Assignees

• Roy Subrata
• Univ Florida

Inventors

• Roy Subrata

Classifications

• B64G1/413

  Ion or plasma engines · CPC title

• F03H1/00

  Using plasma to produce a reactive propulsive thrust (generating plasma H05H1/00 ){(ion sources per se H01J27/02, ion sources for plasma processing or ion beams H01J37/08)} · CPC title

• B64G1/406

  Operations & Transport · mapped topic

• F03H1/0093

  Electro-thermal plasma thrusters, i.e. thrusters heating the particles in a plasma (resistojets per se B64G1/415) · CPC title

• F03H1/0087

  Electro-dynamic thrusters, e.g. pulsed plasma thrusters · CPC title