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Reexamination Certificate
1998-10-22
2001-07-03
Cole, Elizabeth M. (Department: 1771)
Stock material or miscellaneous articles
Web or sheet containing structurally defined element or...
Noninterengaged fiber-containing paper-free web or sheet...
C428S312600, C428S314400, C428S317900, C428S319100, C428S312200
Reexamination Certificate
active
06254975
ABSTRACT:
BACKGROUND AND SUMMARY OF THE INVENTION
This application claims the priority of German Patent Application No. 197 26 598.6, filed Oct. 22, 1997, the disclosure of which is expressly incorporated by reference herein.
The present invention relates to a ceramic composite reinforced by means of C-fibers. Fiber-reinforced ceramic composites have a high temperature stability and, in contrast to monolithic ceramics (i.e., ceramics that are not reinforced by fibers), have a low brittleness (i.e., a high tolerance with respect to damage). Fiber-reinforced ceramic composites are therefore well suitable for use as construction materials for high-temperature components in air and space travel, for example, as heat shields, hot leading wing edges or noses of space transport systems, engines of rockets or components of aircraft engines.
Because of the ratio of stability to density and/or stiffness to density, which is favorable at high temperatures, and because of the possibility of eliminating additional thermal insulation, fiber-reinforced ceramic composites can clearly save weight in comparison to current state of the art ceramics.
So far, known fiber-reinforced ceramic composites are carbon-fiber-reinforced carbon (abbreviated C/C) and carbon-fiber-reinforced silicon carbide (abbreviated C/SiC). Furthermore, composites have been developed that have fibers made of: SiC (for example, having the tradename TYRANNO and NICALON), Al
2
O
3
, the system Al
2
O
3
—SiO
2
, or the system Al
2
O
3
—SiO
2
—B
2
O
3
(for example, having the tradename NEXTEL).
Because of an insufficient thermal stability of the fibers, known ceramic composites with fibers other than carbon C can only be used only at a maximum of 1,100° C. For many applications in air and space travel, this temperature limit represents an exclusion criterion. For example, when space transport systems or engine parts of rockets or airplanes reenter the earth atmosphere, they will clearly have higher temperatures.
On the other hand, all materials containing C-fibers are very sensitive to oxidation. For this reason, C/C can be used at high temperatures only in an oxygen-free atmosphere or in air for a very short time. Even C/SiC is generally not stable with respect to oxidation because (1) C-fibers oxidize from the direction of the edge and, (2) such materials are porous, thus oxygen reaches the C-fibers through the matrix pores and oxidizes the C-fibers.
It is an object of the present invention to provide a construction material particularly for use in air travel and space travel that meets the following requirements:
a) density <2.3 g/cm
3
;
b) applicable from ambient temperature to 1,600° C.;
c) tensile strength above 200 MPa, with simultaneous tolerance to damage, that is, without brittle fracture;
d) stability in an oxidizing atmosphere, particularly air;
e) at least 50 h of service life;
f) resistant to an at least 100-fold thermal shock; and
g) economical manufacturability of complex structures.
No construction material has so far been known that meets all above-mentioned requirements simultaneously. The situation with respect to the individual requirements is as follows:
a) density <2.3 g/cm
3
and
b) applicable from ambient temperature to 1,600° C.
From the combination of these two requirements, it is concluded that the material must contain carbon.
c) Tensile strength above 200 MPa, with simultaneous tolerance to damage; that is, without brittle fracture.
This results in the necessity of having a ceramic composite that is reinforced by C-fibers. As the matrix, the following is conceivable: C, SiC, Si
3
N
4
, SiBCN, SiBN
3
C, or similar materials. The C-fibers may be either short (i.e., chopped or cut fibers having a length of several millimeters to a few centimeters) or endless (i.e., not chopped or uncut and coming from a spool having a length of several hundreds of meters).
d) Stability in an oxidizing atmosphere, particularly air.
For avoiding the oxidation of C-fibers, an oxidation protection system is required. External protective layers are customary, as described, for example, in S. Goujard et al., The oxidation behavior of two- and three-dimensional C/SiC thermostructural materials protected by chemical-vapor-deposition polylayers, 29 J. Mater. Sci. 6212-6220 (1994).
e) More than 50 h of service life in the temperature range 20-1,600° C.
Protective systems that protect the C-fibers in the ceramic composites for more than 50 h in the whole temperature range of 20-1,600° C. from oxidation are not known. Because even improved protective systems would not be sufficient for protecting a carbon matrix against oxidation for this long, only SiC, Si
3
N
4
, SiBNC or similar non-oxides can be used as the matrix. Oxides, such as Al
2
O
3
and ZrO
2
are not compatible with the C-fiber.
f) Resistant to an at least 100-fold thermal shock.
Particularly when combined with the condition that the material must be usable in the whole temperature range of from 20° C. to 1,600° C., this requirement represents a demand that is higher than e). Such material systems are not the state of the art: M. P. Bacos & O. Sudre,
Critical review on oxidation protection for carbon-based composite in High-temperature ceramic-matrix composites
, Ceramic Transactions, Vol. 57 (1995); J. Strife & J. Sheehan,
Ceramic coating for carbon-carbon composites
, Ceramic Bull. 67(2) 369 (1988).
g) Economical manufacturability of complex structures.
For an extremely light construction in air and space travel, integral structures are desired analogous to those that are currently manufactured from C-fiber-reinforced plastic materials. A manufacturing process for C-fiber-reinforced ceramic composites that is derived therefrom and also approaches the final ceramic composite is the polymer infiltration and pyrolysis process as disclosed in U. Trabant et al.,
Test results of low cost C/SiC for martian entry and reusable launcher
, Proceedings 47
th
Intern. Astronautical Cong., China (1996); W. Schaefer et al.,
Hot aerospace structures from fiber reinforced ceramics
, Proceedings 17
th
Conf. Aerospace Material Engineering, Paris (1997); T. Haug et al.,
Herstellverfahren für oxidationsbeständiqe faserverstärkte Keramiken
, Proceedings “Werkstoffwoche '96”, DGM/DKG, Stuttgart (1996). Further background information is available in M. Balat et al.,
Active to passive transition in the oxidation of silicon based ceramics at high temperatures
, Proceedings “3
rd
European Workshop on High Temperature Materials” ESA (1996). It is therefore applicable here.
The object of the present invention is a material that meets the above-mentioned requirements a) to g) and has the following characteristics:
I. Reinforcing fibers made of C;
II. A matrix made of SiC+C (5-30% by weight excess carbon);
III. Pores;
IV. A lower oxidation protection layer having the function of a primer and surface sealing with a coefficient of thermal expansion adapted to the base material, made of one or several materials SiC, SiO
2
, B;
V. A central oxidation protection layer having the function of an oxygen getter by oxide formation and/or local crack sealing by glass formation from one or several Si-containing or B-containing compounds, such as MoSi
2
, Mo
3
Si, Mo
5
Si
3
, SiC, BN, B
4
C, Si
3
N
4
, SiB
6
, TaB
2
, B, SiBCN, SiBN
3
C or similar compounds;
VI. An upper oxidation protection layer having the function of the erosion protection and the adaptation of radiation emission and absorption, made of one or several of the materials SiC, MoSi
2
, Mo
3
Si, Mo
5
Si
3
, ZrB, ZrB
2
, ZrB
12
, Si
3
N
4
, ZrO
2
, Y
2
O
3
, Al
2
O
3
, system Al
2
O
3
—SiO
2
;
VII. The above-mentioned three oxidation protection layers may be composed of partial layers. By means of several partial layers adapted to individual partial temperature intervals, an oxidation protection can be achieved that has a continuously optimal effect from 20° C. to 1,600° C. because individual materials in each case can optimally provide oxidation protection in only limited temperature ranges.
VIII. Manufacturability of the base
Haug Tilman
Hertel Doris
Kayser Ursula
Knabe Helmut
Rebstock Kolja
Cole Elizabeth M.
Crowell & Moring , L.L.P.
Dornier GmbH
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