Hybrid ablative thermal protection systems and associated methods

Patent metadata

Field	Value
Publication number	US-10279931-B2
Application number	US-201615004456-A
Country	US
Kind code	B2
Filing date	Jan 22, 2016
Priority date	Aug 31, 2009
Publication date	May 7, 2019
Grant date	May 7, 2019

Abstract

A method of forming a thermal protection system for high speed aircraft is described. The method includes mechanically working an uncured ablator material into a first surface of a felt layer such that the ablator material penetrates a distance into a thickness of the felt layer thereby forming an region that has a mixture of felt and ablator material, adding additional uncured ablator material to the worked ablator material, and curing the combined ablator material.

First claim

What is claimed is:

1. A thermal protection system for high speed aircraft comprising:
a felt layer comprising a first side, a second side, and a thickness separating said sides; said first side operable for attachment to a structure of the aircraft using an adhesive;
a first layer of silicone-based ablator material mechanically worked into the thickness of said felt layer from said second side prior to the curing of said ablator material, said first layer worked into the thickness of said felt layer at a depth less than the thickness of said felt layer such that the first side is free of silicone-based ablator material; and
a second layer of silicone-based ablator material extending over said second side of said felt layer such that said first and second layers of silicone-based ablator material are combined.

2. A thermal protection system according to claim 1 wherein said first layer is worked into the thickness of said felt layer from said second side to a depth that ranges from about 0.03 inches to about 0.04 inches.

3. A thermal protection system according to claim 1 wherein a thickness of the ablator material and felt layer combination is verified using a needle like device as a depth gauge.

4. A thermal protection system according to claim 1 wherein said felt layer has a thickness of about 0.160 inches.

5. A thermal protection system according to claim 1 wherein said felt layer comprises nomex fibers worked to present a porous surface to said first layer of ablator material prior to curing.

6. A thermal protection system according to claim 1 wherein said first layer of ablator material is milled to a desired thickness after curing.

7. A thermal protection system according to claim 1 wherein said first layer of ablator material is thinned with a solvent prior to being worked into said felt layer, the solvent removed during a curing of said ablator material.

8. A thermal protection system according to claim 1 wherein said ablator material is cured at an ambient temperature for up to 24 hours, the curing temperature subsequently increased at about 10 degrees Fahrenheit per hour until the curing temperature is about 180 degrees Fahrenheit, the curing temperature maintained at approximately 180 degrees Fahrenheit for up to 48 hours.

9. A thermal protection system according to claim 1 , wherein said second layer of silicone-based ablator material is extended over said second side of said felt layer prior to the curing of said first layer of silicone-based ablator material.

10. A thermal protection system according to claim 1 , wherein said felt layer is fabricated from a lower weight material than said ablator material.

11. An aerospace vehicle comprising:
a substructure;
a thermal protection system coupled to said substructure, said thermal protection system comprising:
a felt layer comprising a first side, a second side, and a thickness separating said sides; said first side attached to the substructure using an adhesive;
a first layer of silicone-based ablator material mechanically worked into the thickness of said felt layer from said second side prior to the curing of said ablator material; and
a second layer of silicone-based ablator material extending over said second side of said felt layer such that said first and second layers of silicone-based ablator material are combined.

12. An aerospace vehicle according to claim 11 further comprising a layer of adhesive positioned between said substructure and said felt layer.

13. An aerospace vehicle according to claim 11 , wherein said substructure is an aluminum panel.

14. An aerospace vehicle according to claim 11 , wherein said substructure is an exterior panel of the aerospace vehicle.

15. An aerospace vehicle according to claim 11 , wherein said first layer is worked into the thickness of said felt layer from said second side to a depth that ranges from about 0.03 inches to about 0.04 inches.

16. An aerospace vehicle according to claim 11 , wherein said felt layer has a thickness of about 0.160 inches.

17. An aerospace vehicle according to claim 11 , wherein said felt layer comprises nomex fibers worked to present a porous surface to said amount of ablator material prior to curing.

18. An aerospace vehicle according to claim 11 , wherein said first layer of ablator material is thinned with a solvent prior to being worked into said felt layer, the solvent removed during a curing of said ablator material.

19. An aerospace vehicle according to claim 11 , wherein said second layer of silicone-based ablator material is extended over said second side of said felt layer prior to the curing of said first layer of silicone-based ablator material.

20. An aerospace vehicle according to claim 11 , wherein said felt layer is fabricated from a lower weight material than said ablator material.

Assignees

• Boeing Co

Inventors

• Brewer Bryan Clark
• Mobers Patrick Jacob

Classifications

• B64G1/62

  Systems for re-entry into the earth's atmosphere; Retarding or landing devices · CPC title

• B64G1/58 Primary

  Thermal protection, e.g. heat shields · CPC title

• B64C1/38

  Constructions adapted to reduce effects of aerodynamic or other external heating · CPC title

• B29C65/4835

  Heat curing adhesives · CPC title