U.S. patent application number 09/357282 was filed with the patent office on 2001-12-13 for durable refractory ceramic coating.
Invention is credited to BENNER, KENNETH W., REEVES, SARAH D., WITTENAUER, JEROME P., YASUKAWA, ROBERT D..
Application Number | 20010051218 09/357282 |
Document ID | / |
Family ID | 23404987 |
Filed Date | 2001-12-13 |
United States Patent
Application |
20010051218 |
Kind Code |
A1 |
WITTENAUER, JEROME P. ; et
al. |
December 13, 2001 |
DURABLE REFRACTORY CERAMIC COATING
Abstract
An article used for thermal protection includes a base structure
having at least one surface. The base structure is made of a
ceramic oxide material. A silicide coating is formed on the at
least one surface of the base structure. The coating comprises a
refractory metal and silicon, which together form a silicide. The
coating is at least partially diffused into the base structure at
the at least one surface.
Inventors: |
WITTENAUER, JEROME P.; (PALO
ALTO, CA) ; REEVES, SARAH D.; (SUNNYVALE, CA)
; BENNER, KENNETH W.; (FREMONT, CA) ; YASUKAWA,
ROBERT D.; (SAN JOSE, CA) |
Correspondence
Address: |
EDWARD A. PENNINGTON
SWIDLER BERLIN SHEREFF FRIEDMAN
3000 K STREET, N.W.
WASHINGTON
DC
20007
US
|
Family ID: |
23404987 |
Appl. No.: |
09/357282 |
Filed: |
July 20, 1999 |
Current U.S.
Class: |
427/376.2 ;
427/419.7; 427/427 |
Current CPC
Class: |
B64G 1/226 20130101;
C04B 41/009 20130101; C23C 26/02 20130101; Y02T 50/60 20130101;
Y10T 428/273 20150115; C04B 2111/00982 20130101; F05D 2300/612
20130101; C04B 41/5071 20130101; F05D 2230/90 20130101; F01D 5/288
20130101; B64G 1/58 20130101; C23C 10/26 20130101; Y02T 50/67
20130101; Y02T 50/672 20130101; C04B 41/5071 20130101; C04B 41/4539
20130101; C04B 41/507 20130101; C04B 41/009 20130101; C04B 14/4631
20130101; C04B 30/02 20130101 |
Class at
Publication: |
427/376.2 ;
427/419.7; 427/427 |
International
Class: |
B05D 003/02; B05D
001/36 |
Claims
What is claimed is:
1. A thermal protection system comprising: a base structure having
at least one surface, the base structure being made of a ceramic
oxide material; and a silicide coating formed on the at least one
surface, the coating being at least partially diffused into the
base structure at the at least one surface.
2. A thermal protection system according to claim 1, wherein the
silicide coating is molybdenum silicide.
3. A thermal protection system according to claim 1, wherein the
silicide coating is tantalum silicide.
4. A thermal protection system according to claim 1, wherein the
silicide coating is niobium silicide.
5. A thermal protection system according to claim 1, wherein the
refractory metal is selected from the group consisting of tungsten,
molybdenum, tantalum, niobium, vanadium and chromium.
6. A thermal protection system according to claim 1, wherein the
coating includes boron in an amount sufficient to lower the melting
point of the coating and thereby facilitate fusion of the
coating.
7. A thermal protection system according to claim 1, wherein the
silicide coating includes molybdenum silicide and molybdenum
boride.
8. A thermal protection system according to claim 1, wherein the
base structure is a silica-based rigid fibrous insulation material
of predetermined size and shape.
9. A thermal protection system according to claim 1, wherein the
silicide coating has a weight of between about 0.25 and 2.0 grams
per square inch.
10. A thermal protection system according to claim 1, wherein the
silicide coating is temperature resistant up to about 2,950.degree.
F.
11. A thermal protection system according to claim 1, wherein the
silicide coating has an impact energy for 50% failure of about 0.02
Joules.
12. A method of forming a coated thermal insulation article,
comprising the steps of: forming a base structure from a ceramic
oxide material; forming a slurry from a liquid carrier constituent
and a powdered constituent which contains silicon and a refractory
metal; applying the slurry in a predetermined coating weight to a
surface of the base structure; and heating the coated base
structure for a time and temperature sufficient to fuse the coating
to the base structure.
13. A method according to claim 12, wherein the step of forming a
slurry further includes adding a flux agent to the slurry to allow
low-temperature fusion of the coating and to further enhance the
oxidation resistance of the coating.
14. A method according to claim 13, wherein the flux agent is
selected from the group consisting of molybdenum diboride and boron
nitride.
15. A method according to claim 12, wherein the applying step
includes applying the slurry by a process selected from the group
consisting of spraying and brushing.
16. A method according to claim 12, wherein the heating step
comprises heating the coated base structure for 1-2 hours at a
temperature of between 1,800 and 2,600.degree. F.
17. A method according to claim 13, wherein the powdered
constituent is selected from the group consisting of molybdenum
disilicide, tantalum disilicide, niobium disilicide, molybdenum
diboride, tantalum diboride, niobium diboride, and boron nitride,
and combinations thereof.
18. A product made by the process of claim 12.
19. An article used for thermal protection, comprising: a base
structure having at least one surface, the base structure being
made of a ceramic oxide material; and a silicide coating formed on
the at least one surface, the coating comprising a refractory metal
and silicon, which together form a silicide, the coating being at
least partially diffused into the base structure at the at least
one surface.
20. An article according to claim 19, wherein the silicide coating
is selected from the group consisting of molybdenum silicide,
tantalum silicide, and niobium silicide.
21. An article according to claim 19, further comprising a
diffusion zone disposed between the silicide coating and the base
structure, and being composed of coating material that diffuses
into the base structure, wherein the silicide coating includes
boron in an amount sufficient to lower the melting point of the
coating and thereby facilitate fusion of the coating the base
structure.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention relates generally to refractory
materials, and more specifically, to insulating materials having
refractory ceramic coatings. A base structure is coated with a
refractory silicide coating. The coating is made of a refractory
metal, i.e., those having a melting point above about 1,650.degree.
C., and silicon. The combination of these materials forms a
"silicide" coating.
DESCRIPTION OF THE RELATED ART
[0002] High temperature environments such as those found in
atmospheric reentry, jet turbine combustion, or rocket propulsion,
necessitate the use of thermal protection systems that provide
oxidation protection, high emissivity, and resistance to mechanical
damage. One example of such a system is the tile used to cover the
outer surfaces of the underbody and wings of NASA's Space Shuttle
Orbiter. Each tile is a lightweight, fibrous, silica-based rigid
fibrous insulation unit with a nominal density of nine (9) pounds
per cubic foot (pcf). The tile is made by the assignee of the
present invention and designated as "LI-900.
[0003] A variation of LI-900, called LI-2200, and likewise
manufactured by the assignee herein, is a twenty-two (22) pcf
version of the LI-900 that offers improved strength at a sacrifice
in weight.
[0004] To improve strength, a new class of rigid reusable
insulation was developed. These consist of the following composite
ceramic materials: FRCI, AETB, and HTP. A further variation of
LI-900, made by NASA's Ames Research Center and designated
"FRCI-12," consists of a blend of silica and aluminoborosilicate
fibers. FRCI-12 has a density of twelve (12) pcf. All three of
these materials are currently qualified for use on the Shuttle
Orbiter Fleet.
[0005] "HTP," which stands for "high thermal performance," refers
to a new class of lightweight ceramic material introduced by the
assignee herein around 1982. Basically, this high-strength
insulation is produced by fusing silica and alumina fibers
together. The insulation is produced to a number of standard
densities at a standard composition of twenty-two (22) percent. The
HTP family of insulants has yielded improvements in strength and
maximum temperature capability relative to earlier generations of
ceramic insulation. Also, at about the time HTP was introduced,
NASA (Ames Research Center) introduced the rigid fibrous insulation
material known as "AETB."
[0006] Coatings have been used in conjunction with refractory
metals, such as tantalum, niobium, and molybdenum, to protect the
underlying metallic structures from oxidation and erosion
experienced in high temperature propulsion environments. Silicide
coatings have been used in the past for such purposes.
[0007] The TPS tiles noted above have in the past been protected by
application of reaction cured glass (RCG) coatings. These coatings
were developed in the early 1970's for the LI-900 class of thermal
insulants. RCG is composed mostly of silica with a small amount of
silicon hexa/tetraboride added as a blackening agent and a fluxing
agent. The coating is applied as a 8-12 mils thick layer onto the
surface of a ceramic tile. As a surface coating, RCG has relatively
poor resistance to impact; as a silica-based system, RCG's maximum
temperature capability is limited to its softening point of about
2,700-2,800.degree. F.
[0008] In 1989, NASA's Ames Research Center developed an insulation
product called toughened uni-piece fibrous insulation (TUFI). The
coating was still silica-based, but it contained about twenty (20)
percent molybdenum disilicide as a blackening agent. TUFI products
represented an advancement in the state of the art because the
coating is applied as a surface impregnation, meaning that it
became commingled with the fibers of the insulation tile near the
surface region. The resultant fused coating is a fiber reinforced
glass which is much more durable than the RCG coating. As a
silica-based coating, however, the TUFI product has the same upper
temperature limit as RCG, i.e., 2,700-2,800.degree. F. TUFI has
been successfully applied to FRCI, HTP, and AETB.
[0009] Refractory metal coatings have been used in ceramic
applications. For example, U.S. Pat. No. 5,413,851 to Storer
describes a ceramic carbon fiber coated with a refractory metal or
metal-based ceramic material. The refractory metal materials used
include molybdenum, tantalum, tungsten, niobium, oxides of
aluminum, yttrium, zirconium, hafnium, gadolinium, titanium, erbium
and other rare earth metals. The fibrous materials that are coated
include alumina, alumina-silica, and alumina-boria- silica. The
coatings are used to enhance strength.
[0010] U.S. Pat. No. 4,530,884 to Erickson et al. describes a
ceramic-metal laminate which is used as insulation in high
temperature environments. The composites described therein have an
inner ceramic layer and an outer metal layer and an intermediate
interface layer of a low modulus metallic low density structure.
These composites are principally used as turbine blades of gas
turbine engines.
[0011] U.S. Pat. No. 5,863,846 to Rorabaugh et al. describes a
ceramic insulation used in aerospace applications in which a slurry
is molded from ceramic fiber to form a soft felt mat which is
impregnated with a sol prior to drying. The mat is exposed to a
catalyst that diffuses into the mat and causes the sol to gel.
[0012] U.S. Pat. No. 5,814,397 to Cagliostro et al. describes
ceramic materials used in space re-entry vehicles in which silica
coatings are formed on fibrous insulations. U.S. Pat. No. 5,079,082
to Leiser et al. describes a porous body of fibrous, low density
silica-based insulation material that is impregnated with a
reactive boron oxide-containing borosilicate glass frit, a silicon
tetraboride fluxing agent, and a molybdenum silicide emittance
agent.
[0013] A continuing need exists for improved lightweight thermal
insulation materials that are temperature resistant and physically
durable. In particular, improved coatings, such as those described
below, are needed for all of the advanced rigid, fibrous insulation
materials described above, such as FRCI, HTP, and AETB.
SUMMARY OF THE INVENTION
[0014] An object of the present invention is to provide an
insulative material that exhibits oxidation protection, high
emissivity, and resistance to mechanical damage.
[0015] Another object of the present invention is to provide an
insulative material that is capable of withstanding
high-temperature environment, including those associated with
atmospheric reentry, jet turbine combustion, and rocket
propulsion.
[0016] Still another object of the invention is to provide a
silicide coating that exhibits higher temperature capability than
silica-based coatings.
[0017] These and other objects of the invention are met by
providing a thermal protection system comprising a base structure
having at least one surface, the base structure being made of a
ceramic oxide material, and a coating formed on the at least one
surface, the coating comprising a refractory metal and silicon and
being at least partially infiltrated into the base structure at the
at least one surface.
[0018] The foregoing features and advantages of the present
invention will be further understood upon consideration of the
following detailed description of the invention taken in
conjunction with the accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0019] FIG. 1 is a perspective view of a thermal insulation system
according to the present invention;
[0020] FIG. 2 is an enlarged side elevational view of the thermal
insulation system of FIG. 1;
[0021] FIG. 3 is an photo-micrograph, magnified at 50.times., of
the thermal insulation system of FIG. 1;
[0022] FIG. 4 is a graph showing impact energy vs. failure rate of
the thermal insulation system of FIG. 1, compared to that of a
prior art insulation material; and
[0023] FIG. 5 is a photograph showing two samples of material after
exposure to high temperatures, in which a sample of the present
invention exhibited high temperature resistivity.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
[0024] Referring to FIG. 1, a thermal insulation system is
generally referred to by the numeral 10. The system can be an
insulating structure, such as a tile or any other object shaped and
sized to perform a specific function. The polyhedron shape of the
system 10 was chosen for illustration purposes. Virtually any shape
is contemplated to be within the scope of the present invention. Of
course, a typical system may employ numerous, similarly shaped
objects, such as in the tiles used on the outer surfaces of NASA's
Shuttle Orbiter. For example, the thermal protection system 10
could be a thermal insulation tile having the dimensions of six (6)
inches by six (6) inches by two (2) inches. The tiles may be flat
or profiled to adopt any desirable shape.
[0025] The thermal protection system 10 includes a base structure
12 having at least one outer surface. The base structure 12 is
preferably made of a ceramic oxide material. Preferred ceramic
oxide materials include those advanced examples discussed above.
While other materials can be used, the base structure is preferably
a ceramic oxide, for example: zirconia, alumina, and/or
silica/alumina blends.
[0026] A coating 14 is formed on one or more of the surfaces of the
base structure 12. The coating 14 is made of a refractory metal and
silicon and is at least partially diffused into the base structure
12 at the at least one surface. The refractor silicide coating
employs the same surface impregnation strategy as the silica-based
TUFI materials used in the past, but substitutes a composition that
is nearly 100% refractory metal silicide. The coating 14 is formed
from any of the refractory metals, meaning those elements with
melting points in excess of 1,650.degree. C., in combination with
silicon. This combination produces a "silicide" coating. Examples
include molybdenum silicide, tantalum silicide and niobium
silicide.
[0027] A particularly preferred material for the coating 14 is
molybdenum silicide. Molybdenum disilicide has a melting point that
is approximately 350.degree. F. higher than silica. The result is
that the silicide coating has higher temperature capability than
the silica-based coatings used in RCG and TUFI materials. Moreover,
since the silicide coatings described herein are applied as a
surface impregnation, the present materials are as durable as the
TUFI materials. RCG and TUFI silica-based coatings have a melting
point of about 3,180.degree. F. and a softening point of between
about 2,700 and 2,800.degree. F. In contrast, the present
silicide-based coatings have a melting point of about 3,540.degree.
F., and a softening point between about 2,900 and 3,000.degree.
F.
[0028] Any of the refractory metals can be used as a constituent of
the coating 14, including molybdenum, tantalum, niobium, vanadium,
chromium, and tungsten, or any other materials that are comparable
chemically and with respect to melting point.
[0029] The molybdenum silicide coating 14 is prepared from high
purity molybdenum disilicide and molybdenum diboride powders. Boron
nitride and other forms of boron may be used as an alternative
source of boron. The silicide provides the desired high melting
point and high emissivity characteristics. The boride serves as a
flux to allow low temperature fusion of the coating and also plays
a role in providing oxidation resistance to the coating. As with
RCG and TUFI coatings, these powders are formulated into a slurry
and applied at a pre-specified coating weight to the surface of the
ceramic oxide base structure 12 by either brushing or spraying. The
coating is fused in an air furnace at a temperature in the range of
about 1,800.degree. F. to about 2,600.degree. F. A more preferred
range is between 2,100.degree. F. and 2,400.degree. F.
[0030] The slurry which contains the constituent powdered materials
includes a liquid carrier, which is a polymeric stock solution
known and used in the art. The heating time in the air furnace can
be anytime sufficient to allow fusion and attachment of the coating
to the base structure. A typical time range is between one (1) and
two (2) hours. The temperature is likewise selected to effect
fusion of the coating, and should be kept below the sintering
temperature of the substrate, or base structure, which for the
aforementioned preferred materials is about 2,600.degree. F.
[0031] The coating 14 can be varied in thickness. "Thickness"
refers to the coating material on the surface and that which has
diffused into the base material, forming an infiltration zone
beneath the surface. In a particularly preferred embodiment, for
tiles used on the X-33 spacecraft currently under development, the
coating 14 has a weight of about 1.0 gram per square inch, which
provides a coating of about 0.030 inches. In this particular
structure, the thickness of 0.030 inches represents only the
surface coating and not the infiltration zone. In general, the
infiltration zone varies depending on the substrate material. A
generally preferred range of weights is 0.25 to 2.0 grams per
square inch for the surface coating (excluding the infiltration
zone). The coating durability tends to increase with higher coating
weights. It will be recognized by those skilled in the art that a
broader range of weights and thickness are possible, depending on
the specifications of the end product.
[0032] The coating 14 is applied to achieve a true surface
impregnation resulting in a silicide coating that is reinforced by
the fiber network of the underlying silica-based rigid fibrous
insulation material which forms the base structure 12. this can be
seen in FIG. 3, which is a photomicrograph of a base structure
coated with molybdenum disilicide. The molybdenum disilcide is
shown as the lighter colored regions, and the ceramic oxide
material which forms the base structure is shown as the darker
regions. It is evident that the molybdenum disilicide has diffused
into the base structure. The infiltration zone, referred to above,
corresponds to the distance between the outer surface of the
surface coating and the lower-most area of penetration of the
coating material into the base material. Referring to FIG. 3, the
thickness of the surface is approximately shown by the letter "A"
and the infiltration zone, which includes the surface depth, is
shown by the letter "B." The result is a highly durable
coating.
[0033] Durability of products made according to the present
invention is demonstrated with reference to FIG. 4, which shows
impact test results for LI-900 prior art coated ceramic thermal
protection tiles, and HTP-6 which incorporates the refractory
silicide coatings of the present invention. The test was conducted
as follows: A steel ball having a volume of one (1) cubic
centimeter was dropped from a fixed height and impacted the surface
with an energy measured in Joules (along the x-axis of FIG. 4). For
each impact energy, five drop tests were conducted and the fraction
of coating failures was noted. The drop height for 0.002J was 2.5
cm; for 0.007 the drop height was 8.0 cm; for 0.026J the drop
height was 32.0 cm; and for 0.042 the drop height was 51.0 cm.
Coating failure was defined as follows: "damage threshold" for RCG
coating is the appearance of a surface crack; "Brinnell failure"
for RCG coating is large area cracking of the RCGT accompanied by
crushing of underlying LI-900 material; and "damage threshold" for
LMMS refractory silicide coating is surface spall. For the latter,
the coating does not fail by cracking as does RCG. As is evident
from the plotted points, the present invention achieved a factor of
25 improvement in impact energy.
EXAMPLE
[0034] To make tiles used for the X-33 spacecraft, a slurry was
created by using 43% by weight molybdenum silicide powder (100
mesh), 7% by weight boron nitride powder (325 mesh), 2.5% by weight
liquid stock solution A, which is a polymer-based binder, and 47.5%
denatured ethanol. These materials are combined and jar milled to
produce a spray slurry which has a density of 11.5+/-0.5 pounds per
gallon, and a particle size distribution as follows:
1 Mass % finer Microns 90 5.25 to 13.76 50 3.30 to 5.57 10 .51 to
1.50
[0035] The slurry is then sprayed to desired thickness and then
dried at temperatures and times specified above.
[0036] Products made according to the present invention were tested
against the prior art and were shown to have marked increases in
thermal performance and durability. In a comparative burner rig
test, refractory silicide coatings of the present invention were
tested against a NASA Ames TUFI+RCG. Both the present coatings and
the NASA coatings were made on a base structure made of HTP 8-22.
The TUFI+RCG coating had a weight of 1.07 g/in.sup.2 and the
present coating, a molybdenum silicide (MoSi.sub.2+MoB.sub.2), had
a total coating weight of 0.57 g/in.sup.2.
[0037] Test samples were exposed to an array of torches fueled by
acetylene, hydrogen, and oxygen. The temperature was measured by
optical pyrometer and thermocouple. FIG. 5 shows two test samples,
one being the present invention and the other the prior art, after
6 minutes of heat at 1590.degree. C. The sample on the right shows
heavy pitting, while the present invention, the sample on the left,
survived the high temperature heating. Other tests have shown that
the present coatings can withstand temperatures in a flame test of
up to about 2,950.degree. F.
[0038] While advantageous embodiments have been chosen to
illustrate the invention, it will be understood by those skilled in
the art that various changes and modifications can be made therein
without departing from the scope of the invention as defined in the
appended claims.
* * * * *