U.S. patent application number 11/728620 was filed with the patent office on 2007-09-27 for low-density ablative heat shield fabrication.
This patent application is currently assigned to ELORET Corporation. Invention is credited to M. Alan Covington, Margaret M. Stackpoole.
Application Number | 20070224407 11/728620 |
Document ID | / |
Family ID | 38533815 |
Filed Date | 2007-09-27 |
United States Patent
Application |
20070224407 |
Kind Code |
A1 |
Covington; M. Alan ; et
al. |
September 27, 2007 |
Low-density ablative heat shield fabrication
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; M. Alan; (San
Jose, CA) ; Stackpoole; Margaret M.; (Santa Clara,
CA) |
Correspondence
Address: |
John P. Wooldridge, Esq.
114 Honu'ea Pl
Kihei
HI
96753
US
|
Assignee: |
ELORET Corporation
|
Family ID: |
38533815 |
Appl. No.: |
11/728620 |
Filed: |
March 26, 2007 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
60785930 |
Mar 24, 2006 |
|
|
|
Current U.S.
Class: |
428/292.1 |
Current CPC
Class: |
B32B 2260/046 20130101;
B32B 5/28 20130101; B32B 2307/304 20130101; B32B 2605/18 20130101;
B64G 1/58 20130101; B32B 27/42 20130101; B32B 27/28 20130101; B32B
27/38 20130101; Y10T 428/249924 20150401; B32B 2262/106 20130101;
B32B 2262/105 20130101; B32B 5/26 20130101; B32B 2260/023
20130101 |
Class at
Publication: |
428/292.1 |
International
Class: |
D04H 13/00 20060101
D04H013/00 |
Claims
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.degree. C. to
300.degree. 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
[0001] 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.
[0002] 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
[0003] 1. Field of the Invention
[0004] 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.
[0005] 2. Description of Related Art
[0006] 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.
[0007] 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.
[0008] 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.
[0009] 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 (.about.12.4 km/sec,
.about.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/cm.sup.2. 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.
[0010] 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
[0011] It is an object of the present invention to provide methods
for making ablative heat shields.
[0012] It is another object to provide a solid, one-piece,
monolithic ablative heat shield.
[0013] 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.
[0014] These and other objects will be apparent based on the
disclosure herein.
[0015] 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
[0016] 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.
[0017] FIG. 1 shows an exemplary individual refractory fiber matrix
or refractory porous substrate block.
[0018] FIG. 2 illustrates the impregnation of a refractory fiber
matrix or refractory porous substrate block with resin from all
sides.
[0019] 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.
[0020] FIG. 4 shows an oven containing an exemplary heat shield
assembly.
[0021] FIG. 5 shows a one-piece bonded heat shield attached to a
spacecraft structure.
[0022] 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.
[0023] FIG. 6B shows a front view if the PICA material of the test
apparatus of FIG. 6A.
DETAILED DESCRIPTION OF THE INVENTION
[0024] 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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.
[0029] Each block is then heat treated in an autoclave or oven to
provide partial cross-linking of the polymer impregnant throughout
the block.
[0030] The sides 16 and 18 and the bottom surface 14 of each block
are then machined to required finish dimensions.
[0031] 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.
[0032] 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.
[0033] 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.
[0034] 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.
[0035] 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.
[0036] 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.
[0037] 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.TM.) and a phenolic resin (SC1008.TM.) 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.
[0038] 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/cm.sup.2 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.
[0039] 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.
* * * * *