SmallSat hybrid propulsion system

Patent metadata

Field	Value
Publication number	US-11578682-B2
Application number	US-201916455532-A
Country	US
Kind code	B2
Filing date	Jun 27, 2019
Priority date	Jun 29, 2018
Publication date	Feb 14, 2023
Grant date	Feb 14, 2023

Abstract

A hybrid propulsion system for a small satellite package consisting of a main rocket motor containing a solid propellant with multiple oxidizer tanks positioned to direct oxidizer into the rocker motor, thereby producing a desired thrust necessary for orbit insertion and/or orbit correction. Additionally, oxidizers can serve a dual function in controlling cold fuel thrusters for attitude adjustment.

First claim

What is claimed is:

1. A CubeSat propulsion system comprising:
a CubeSat form factor having a defined single volume;
a main propulsion motor vessel centrally disposed within the defined single volume and having a body with a forward end and an aft end and being formed of an outer wall and an inner wall, wherein the inner wall forms a central cavity, and wherein the inner wall is lined with an insulative material that forms a thermal protection layer between the inner wall and a solid rocket fuel, the solid rocket fuel being disposed within the central cavity and bonded to the insulative material such that the solid rocket fuel is positioned where a majority of the central cavity is filled with the solid rocket fuel,
wherein the aft end further comprises a flight thrust nozzle having a nozzle throat section and a nozzle exit section wherein the nozzle throat section is positioned near the aft end and the nozzle exit section is disposed distal to the aft end,
multiple oxidizer containment vessels disposed within the defined single volume and dispersed around the main propulsion motor vessel and wherein each oxidizer vessel of the multiple oxidizer vessels is fluidly connected to the main propulsion motor vessel, such that each oxidizer vessel of the multiple oxidizer vessels delivers an oxidizer upstream of the solid rocket fuel, and into the main propulsion motor vessel within a predefined channel disposed within the solid rocket fuel, and
an ignition source mechanically connected to the main propulsion motor vessel wherein the ignition source operates to vaporize a portion of the solid rocket fuel to generate a vaporized fuel and wherein the oxidizer delivered to the main propulsion motor vessel interacts with the vaporized fuel to produce combustion along the length of the predefined channel and produce an exhaust thrust through the flight thrust nozzle such that a solid rocket fuel utilization rate of 97% or higher is achieved due to the bonding of the solid rocket fuel and the insulative material.

2. The CubeSat propulsion system of claim 1 , wherein the solid rocket fuel is a Poly Methyl Methacrylate (PMMA).

3. The CubeSat propulsion system of claim 2 , wherein the solid rocket fuel is a clear PMMA.

4. The CubeSat propulsion system of claim 1 , wherein the multiple oxidizer containment vessels comprises four oxidizer vessels.

5. The CubeSat propulsion system of claim 1 , wherein each oxidizer containment vessel of the multiple oxidizer containment vessels comprises a respective composite overwrap vessel having a respective lightweight metallic liner with a respective composite overwrap.

6. The CubeSat propulsion system of claim 5 , wherein each metallic liner is selected from a group consisting of: steel, an alloy comprising a majority by weight of nickel and copper, and an alloy comprising a majority by weight of nickel and chromium.

7. The CubeSat propulsion system of claim 5 , wherein the composite overwrap is made from a material selected from a group consisting of carbon and Poly-paraphenylene terephthalamide.

8. The CubeSat propulsion system of claim 1 wherein the CubeSat form factor is a 6 U configuration.

9. The CubeSat propulsion system of claim 1 wherein the CubeSat form factor is a 12 U configuration.

10. The CubeSat propulsion system of claim 2 , wherein the solid rocket fuel is a black PMMA.

11. The CubeSat propulsion system of claim 1 , wherein the insulative material is selected from a group consisting of a rubber type material and carbon phenolic.

12. The CubeSat propulsion system of claim 1 , further comprising a plurality of cold gas thrusters in fluid communication with the multiple oxidizer containment vessels such that the plurality of cold gas thrusters are capable of producing thrust for attitude control.

13. The CubeSat propulsion system of claim 12 , wherein the cold plurality of gas thrusters are selected from a group consisting of Thrust Vector Control thrusters and Attitude Control thrusters.

14. The CubeSat propulsion system of claim 1 , wherein the ignition source is selected from a group consisting of pyrotechnic ignitors, augmented spark ignitors and laser ignitors.

15. The CubeSat propulsion system of claim 1 , wherein the solid rocket fuel further comprises a centralized burn channel the runs from the forward end the aft end.

16. The CubeSat propulsion system of claim 15 , where the centralized burn channel is a single centralized channel.

17. The CubeSat propulsion system of claim 15 , wherein the centralized burn channel is a preformed configuration of multiple channels.

18. A SmallSat propulsion system comprising:
a SmallSat volume having a mass of less than 200 kg;
a main propulsion motor vessel centrally disposed and having a body with a forward end and an aft end and being formed of an outer wall and an inner wall, wherein the inner wall forms a central cavity, and wherein the inner wall is lined with an insulative material that forms a thermal protection layer between the inner wall and a solid rocket fuel, the solid rocket fuel being disposed within the central cavity and bonded to the insulative material such that the solid rocket fuel is positioned where a majority of the central cavity is filled with the solid rocket fuel,
wherein the aft end further comprises a flight thrust nozzle having a nozzle throat section and a nozzle exit section wherein the nozzle throat section is positioned near the aft end and the nozzle exit section is disposed distal to the aft end,
multiple oxidizer containment vessels dispersed around the main propulsion motor vessel and wherein each oxidizer vessel of the multiple oxidizer vessels is fluidly connected to the main propulsion motor vessel, such that each oxidizer vessel of the multiple oxidizer vessels delivers an oxidizer upstream of the solid rocket fuel, and into the main propulsion motor vessel within a predefined channel disposed within the solid rocket fuel, and
an ignition source mechanically connected to the main propulsion motor vessel wherein the ignition source operates to vaporize a portion of the solid rocket fuel to generate a vaporized fuel and wherein the oxidizer delivered to the main propulsion motor interacts with the vaporized fuel to produce combustion along a length of the predefined channel and produce an exhaust thrust through the flight thrust nozzle such that a solid rocket fuel utilization rate of 97% or higher is achieved due to the bonding of the solid rocket fuel and the insulative material.

19. The SmallSat propulsion system of claim 18 , wherein the solid rocket fuel is a Poly Methyl Methacrylate (PMMA).

20. The SmallSat propulsion system of claim 19 , wherein the solid rocket fuel is a clear PMMA.

21. The SmallSat propulsion system of claim 18 , the multiple oxidizer containment vessels comprise two oxidizer vessels.

22. The SmallSat propulsion system of claim 18 , wherein the multiple oxidizer containment vessels are a composite overwrap vessel having a lightweight metallic liner with a composite overwrap.

23. The SmallSat propulsion system of claim 22 , wherein the metallic liner is selected from a group consisting of: steel, an alloy comprising a majority by weight of nickel and copper, and an alloy comprising a majority by weight of nickel and chromium.

24. The SmallSat propulsion system of claim 22 , wherein the composite overwrap is made from a material selected from a group consisting of carbon and Poly-paraphenylene t

Assignees

• California Inst Of Techn

Inventors

• Karp Ashley C
• Jens Elizabeth
• Rabinovitch Jason
• Nakazono Barry
• Conte Antonietta

Classifications

• F02K9/80

  characterised by thrust or thrust vector control (burning control of solid propellants F02K9/26; feeding control of liquid or gaseous propellants F02K9/56; re-ignitable, restartable or intermittently operated rocket-engine plants F02K9/94) · CPC title

• F02K9/32 Primary

  Constructional parts; Details not otherwise provided for · CPC title

• F02K9/95

  characterised by starting or ignition means or arrangements (safety devices F02K9/38) · CPC title