Title:
Low-density ablative heat shield fabrication
Kind Code:
A1


Abstract:
Spacecraft heat shields are fabricated as one-piece assemblies using low-density ablative thermal protection materials. The heat shield assembly is built from modular pieces formed by ablative impregnation processing. Once the full-size heat shield is assembled from the modular blocks, heat treatment is used to bond the individual blocks together by facilitating polymeric cross-linking of impregnant material within and/or between each block. This provides a structurally integral one-piece heat shield assembly that can be further machined to final dimensions and attached directly to a spacecraft structure or a carrier panel separately attached to the spacecraft



Inventors:
Covington, Alan M. (San Jose, CA, US)
Stackpoole, Margaret M. (Santa Clara, CA, US)
Application Number:
11/728620
Publication Date:
09/27/2007
Filing Date:
03/26/2007
Assignee:
ELORET Corporation
Primary Class:
International Classes:
D04H13/00
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Primary Examiner:
O'HARA, BRIAN M
Attorney, Agent or Firm:
John P. Wooldridge, Esq. (Kihei, HI, US)
Claims:
We claim:

1. A method for fabricating an ablative heat shield, comprising: impregnating each piece of a plurality of pieces of ablative thermal protection material with polymeric resin; assembling said plurality of pieces into a desired heat shield shape; and heat-treating said shape to bond said plurality of pieces into a solid heat shield.

2. The method of claim 1, wherein said ablative thermal protection material comprises fiber matrix material.

3. The method of claim 1, wherein said ablative thermal protection material comprises refractory porous substrate material.

4. The method of claim 1, wherein each said piece of said plurality of pieces is selected from a group consisting of a modular piece and a block.

5. The method of claim 1, wherein each said piece is machined to press fit to another said piece.

6. The method of claim 2, wherein said fiber matrix material comprises refractory fiber matrix material.

7. The method of claim 1, wherein said ablative thermal protection material is selected from the group consisting of carbon, silica, alumina, aluminosilicate and silicon carbide.

8. The method of claim 1, wherein said polymeric resin comprises ablative thermal protection materials.

9. The method of claim 1, wherein said polymeric resin comprises phenolic resin.

10. The method of claim 9, wherein said phenolic resin comprises Phenolic Impregnated Carbon Ablator.

11. The method of claim 1, wherein said polymeric resin comprises a phenolic, silicone, epoxy or pre-ceramic polymer compound.

12. The method of claim 1, wherein the step of heat-treating comprises polymeric cross-linking (i) within and between each said piece or (ii) within or between each said piece.

13. The method of claim 1, further comprising machining said solid heat shield to desired dimensions.

14. The method of claim 13, further comprising attaching said heat shield to a spacecraft structure.

15. The method of claim 13, further comprising attaching said heat shield to a carrier panel.

16. The method of claim 15, further comprising attaching said heat carrier panel to a spacecraft structure.

17. The method of claim 1, where the step of impregnating each piece of a plurality of pieces fiber matrix material with polymeric resin comprises immersing said plurality of pieces into said polymeric resin.

18. The method of claim 1, further comprising remove excess solvent from said heat shield.

19. The method of claim 1, wherein the step of assembling said plurality of pieces into a desired heat shield shape includes assembling said pieces against a mandrel.

20. The method of claim 1, wherein the step of heat-treating is carried out at a temperature within a range from 50° C. to 300° C.

21. An ablative heat shield, comprising: a plurality of pieces of ablative thermal protection material configured in a heat shield form; polymeric resin attached to said ablative thermal protection material; and a bond between adjacent pieces of said plurality of pieces, wherein said bond comprises polymeric cross-linking.

22. The apparatus of claim 21, wherein said ablative thermal protection material comprises fiber matrix material.

23. The apparatus of claim 21, wherein said ablative thermal protection material comprises refractory porous substrate material

24. The apparatus of claim 21, wherein each said piece of said plurality of pieces is selected from the group consisting of a modular piece and a block.

25. The apparatus of claim 22, wherein said fiber matrix material comprises refractory fiber matrix material.

26. The apparatus of claim 25, wherein said refractory fiber matrix material is selected from the group consisting of carbon, silica, alumina, aluminosilicate and silicon carbide.

27. The apparatus of claim 21, wherein said polymeric resin comprises ablative thermal protection materials.

28. The apparatus of claim 21, wherein said polymeric resin comprises phenolic resin.

29. The apparatus of claim 28, wherein said phenolic resin comprises Phenolic Impregnated Carbon Ablator.

30. The apparatus of claim 21, wherein said polymeric resin comprises a phenolic, silicone, epoxy or pre-ceramic polymer compound.

Description:

This application claims priority to U.S. Provisional Patent Application Ser. No. 60/785,930, titled “Low-Density Ablative Heat Shield Fabrication,” filed Mar. 24, 2006 and incorporated herein by reference.

The invention described herein was made by nongovernment employees, whose contributions were done in the performance of work under a NASA contract, and is subject to the provisions of Public Law 96-517 (35 U.S.C. 202). This invention was made with Government support under contract NNA04BC25C awarded by NASA. The Government has certain rights in this invention.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to the space program, and more specifically, it relates to processes for making ablative heat shields that provide protection to spacecraft during the severe heating conditions of atmospheric entry.

2. Description of Related Art

The making of large ablative heat shields that provide protection to spacecraft during the severe heating conditions of atmospheric entry is a difficult problem. Limitations due to inherent physical properties and the inability to fabricate large assemblies due to processing and/or machining requirements have prevented the use of the most promising materials in spacecraft design in several instances. In the specific cases of low density and mid-density ablative materials, difficulties in controlling the impregnation of fiber matrix materials with polymeric resins, and difficulty in processing larger fiber matrix substrates within density specifications to achieve uniform material properties have limited the size that billets or blocks of these materials can be made. These limitations have required the use of less capable materials or the use of fabrication and assembly methods that lead to complex and costly final products. A new fabrication process is desired that would overcome several of the problems previously encountered in the making of large, one-piece heat shields. Such process should be applicable to a wide variety of refractory fiber matrix materials or refractory porous substrates and polymeric resins that are known to form efficient ablative heat shield materials.

The type of thermal protection system (TPS) that best protects against high heat flux is the ablative heat shield. The ablative heat shield functions by the energy-absorbing thermal degradation of a polymeric component resulting in the production of a char layer and of gaseous products through a process known as pyrolysis; the absorption of additional energy as these gases flow through the porous degraded material to the heat shield surface; the possible phase change of components from solid to liquid to gaseous, or from solid to gaseous states; and the reduction of the convective heat flux to the heat shield surface by gaseous products as they leave the surface by the thickening and cooling of the boundary layer in a process called blowing. The kinetics and products of the pyrolysis process can be measured in real time using thermogravimetric analysis, so that the ablative performance can be evaluated. Ablation can also provide blockage against radiative heat flux by the introduction of spectrally absorbing gaseous pyrolysis products into the boundary and shock layers in front of an entry spacecraft. Radiative heat flux blockage was the primary thermal protection mechanism of the Galileo Probe TPS material (carbon phenolic). Thermal protection also can be enhanced in some TPS materials through coking. Coking is the process of forming and depositing solid carbon within the char layer of the TPS, resulting in a localized density increase within the char.

The thermal conductivity of a TPS material is proportional to the material's density and dependent on fiber orientation in fiberous substrates. Carbon phenolic is a very effective ablative material but also has high density and resulting high conductivity which is undesirable. If the heat flux experienced by an entry vehicle is insufficient to cause pyrolysis then the TPS material's conductivity could allow heat flux conduction into the TPS attachment and spacecraft structures, thus leading to TPS failure. Consequently for entry trajectories causing lower heat flux, higher density TPS materials such as carbon phenolic are inappropriate and lower density TPS materials may be better design choices.

The concepts presented herein are applicable to a broad range of ablative TPS constructed from refractory fibrous matrix materials or refractory porous substrates and polymeric impregnation resins. An example used herein for purposes of illustration is Phenolic Impregnated Carbon Ablator (PICA) TPS. This material was developed by NASA Ames Research Center and was the primary TPS material for the Stardust Sample Return Capsule aeroshell. Because the Stardust spacecraft was the fastest man-made object to reenter Earth's atmosphere (˜12.4 km/sec, ˜28,000 mph relative velocity at 135 km altitude), PICA was an enabling technology for the Stardust mission. (For reference, the Stardust reentry was faster than the Apollo Mission capsules and 70% faster than the reentry velocity of the Shuttle.) PICA is a modern TPS material that has the advantages of low density (much lighter than carbon phenolic) coupled with efficient ablative capability at high heat flux. Stardust's heat shield (0.81 m base diameter) was manufactured from a single monolithic piece sized to withstand a nominal peak heating rate of 1200 W/cm2. PICA is a good choice for ablative applications such as high-peak-heating conditions found on sample-return missions or lunar-return missions. PICA's thermal conductivity is lower than other high-heat-flux ablative materials, such as conventional carbon phenolics.

Small blocks or tiles of both ablative and insulative thermal protection materials have been used in heat shield applications (e.g., Space shuttle Orbiter) because of thermal expansion and processing size limitations. This requires that each tile be machined to a finished size prior to being individually attached to spacecraft structure in a complex and costly operation and, in the case of the Space Shuttle, fabric filler materials were required to fill gaps between tiles that exist before entry heating. In other applications (e.g., Genesis sample return capsule) the choice of an ablative heat shield material was constrained by the inability to make the required one-piece heat shield. For the Apollo capsule heat shield, the required one-piece ablative heat shield was fabricated by injecting an epoxy resin into a phenolic honeycomb in a costly, complex, and hard to control process.

SUMMARY OF THE INVENTION

It is an object of the present invention to provide methods for making ablative heat shields.

It is another object to provide a solid, one-piece, monolithic ablative heat shield.

Sill another object is to provide an ablative heat shield of modular pieces of a fiber matrix material or refractory porous substrate material that has been cross-linked together to form a one-piece assembly.

These and other objects will be apparent based on the disclosure herein.

The fabrication of large (larger than 1 meter diameter) spacecraft heat shields as one-piece assemblies using low-density ablative thermal protection materials (TPS) formed from rigid substrates has been constrained by limits on available component matrix material sizes and processing requirements. Methods are provided that allow large uni-piece heat shields to be fabricated for use on future space vehicles that require protection from atmospheric entry heating at severe conditions. These fabrication methods provide such large assemblies by building a heat shield assembly from modular pieces (blocks) formed by conventional ablative TPS impregnation and processing methods and limitations. Once the full-size heat shield is assembled from the modular blocks, appropriate heat treatment is used to bond the individual blocks together by facilitating polymeric cross-linking of impregnant material within and/or between each block. This provides a structurally integral one-piece heat shield assembly that can be further machined to final dimensions and attached directly to spacecraft structure or a carrier panel separately attached to the spacecraft.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings, which are incorporated into and form a part of the disclosure, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

FIG. 1 shows an exemplary individual refractory fiber matrix or refractory porous substrate block.

FIG. 2 illustrates the impregnation of a refractory fiber matrix or refractory porous substrate block with resin from all sides.

FIG. 3 shows the assembly on a mandrel of a section of a heat shield having a plurality of fiber matrix or refractory porous substrate blocks.

FIG. 4 shows an oven containing an exemplary heat shield assembly.

FIG. 5 shows a one-piece bonded heat shield attached to a spacecraft structure.

FIG. 6A shows a side-view of a test sample of bonded PICA attached to a low density ceramic fiber insulation which is attached to a metal model holder.

FIG. 6B shows a front view if the PICA material of the test apparatus of FIG. 6A.

DETAILED DESCRIPTION OF THE INVENTION

A new fabrication process is proposed that overcomes several of the problems previously encountered in the making of large, one-piece heat shields. This process is applicable to any ablative material formed from a low density, high-temperature rigid fiber matrix or low density refractory porous substrate and impregnated with a polymeric substance that can be chemically cross-linked. Some common known forms of such refractory fiber matrix materials and refractory porous substrates are those from carbon, silica, alumina, aluminosilicates, and silicon carbides. Common polymeric resins used in ablative materials are based on phenolic, silicone, and epoxy compounds as well as some pre-ceramic polymer precursors to silica and silicon oxycarbide systems.

The impregnation process typically involves the immersion of the rigid fiber matrix into the liquid resin solution so that the resin fills a specified and controlled fraction of the void volume space in the fiber matrix or refractory porous substrate to form an uncross-linked material having specified mechanical and thermal properties as an ablative heat shield. After drying to remove excess solvent, the material is processed through a specific heat cycle to partially cross-link the polymer resin to create an efficient ablative material with tailored properties that can be further processed to form large scale heat shield structures. An exemplary process is described below.

FIG. 1 shows an exemplary individual fiber matrix or refractory porous substrate block 10. The top surface 12, the bottom surface 14, and the sides 16 and 18, of the individual fiber matrix or refractory porous substrate blocks are machined to oversize dimensions.

The fiber matrix or refractory porous substrate blocks are impregnated with a selected ablative polymer resin. FIG. 2 illustrates the impregnation from all sides, as indicated by arrows 20, of fiber matrix block 10 with the ablative polymer resin.

The excess impregnant is then cleaned from the sides of the fiber matrix blocks, which now contains uncross-linked resin. Steps 1-3 are carried out for a number of fiber matrix blocks.

Each block is then heat treated in an autoclave or oven to provide partial cross-linking of the polymer impregnant throughout the block.

The sides 16 and 18 and the bottom surface 14 of each block are then machined to required finish dimensions.

FIG. 3 shows the assembly on a mandrel of a section of a heat shield having a plurality of fiber matrix or refractory porous blocks. The figure shows a partial side view of blocks 22 and 30 and shows a full side view of blocks 24, 26 and 28 all on a mandrel 32. Arrows 34 and 36 are used in the figure to illustrate the use of compression to form a full-sized, one-piece heat shield assembly in the desired configuration. Additional polymeric resin of the same or different chemical composition as the impregnant is applied to mating surfaces of each block to enhance the final joint strength and characteristics.

FIG. 4 shows an oven 40 containing an exemplary heat shield assembly 42 having an exemplary shape formed by the previously described steps. Heat shield assembly 42, still on a mandrel 44, is cured in an autoclave or oven 40 as required for partial cross-linking of the polymer between blocks to form an integral, bonded one-piece heat shield without gaps.

The heat shield assembly is then separated from the inner mandrel and the inner and outer mold-line surfaces are machined to the required finish contour and dimensions.

The heat shield assembly is then attached to a carrier panel or to spacecraft structure as required. This attachment may use conventional methods such as high temperature adhesive bonding or mechanical mechanisms. FIG. 5 shows a one-piece bonded heat shield 50 attached to a spacecraft structure 52.

The heat shield produced by the method described above uses separate steps of impregnating and curing each material block with polymer resin, and machining the mating sides and inner surface before forming a larger assembly. An alternate technique is to take advantage of the inherent “glue” of the ablative polymer itself that is contained in the matrix material. Ablative material blocks are assembled without curing and machining, thus eliminating steps described in [0026], [0027], and [0028] to form large integral heat shields. The heat shield assembly then can be cured and finish machined to its final size and directly attached to a spacecraft structure or a carrier panel as a one-piece assembly.

It is expected that the fabrication technique described herein would be widely used for the construction of large spacecraft heat shields made from low to mid-density ablative materials as well as in other applications where module sizes are limited by raw material or processing requirements. Other potential applications are the construction of large-scale thermal protective shielding for containment of high energy devices such as laser test facilities or electrical discharge facilities where one-piece assemblies have advantages in performance and installation for installation and performance. The technique also can be used to form one piece assemblies of high temperature materials used for thermal insulation of containment walls (e.g., furnaces and ovens) by employing a suitable resin to attach together individual un-impregnated fiber matrix blocks for easy installation. Independent mechanical attachments or restraints would be used in this case to retain the blocks in a desired configuration after the resin has degraded into gaseous products and no longer provides bonding.

The concept has been demonstrated in small-scale assemblies and tests. Work to date has concentrated on making and testing such assemblies using a commercially available carbon fiber matrix material (Fiberform™) and a phenolic resin (SC1008™) that was been processed into a promising spacecraft heat shield material called Phenolic Impregnated Carbon Ablator (PICA). FIG. 6A shows a side-view of a test sample of bonded PICA 60 attached to a low density ceramic fiber insulation 62, that is attached to a metal model holder 64. FIG. 6B shows a front view if the PICA material of the test apparatus of FIG. 6A.

Small (3 inch diameter) flat-faced models have been made of this bonded material and tested in an electric discharge arc-jet facility in air at a convective heating flux of about 700 W/cm2 and a pressure of about 0.5 atm to simulate Earth atmospheric entry. The bonded material performed as expected without any failures of the material or the joints, and has established the feasibility of fabricating large heat shield structures by polymer bonding of modular blocks of material. Work is progressing on making larger assemblies and further arc-jet tests are planned.

The foregoing description of the invention has been presented for purposes of illustration and description and is not intended to be exhaustive or to limit the invention to the precise form disclosed. Many modifications and variations are possible in light of the above teaching. The embodiments disclosed were meant only to explain the principles of the invention and its practical application to thereby enable others skilled in the art to best use the invention in various embodiments and with various modifications suited to the particular use contemplated. The scope of the invention is to be defined by the following claims.