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Materials for advanced space propulsion systems
Materials Science and Engineering, A143 ( 1991 ) 2 1 - 2 9
21
Materials for advanced space propulsion systems Neil E. Paton Rockwell International Corporation, Roeketdyne Division, 6633 Canoga Avenue, Canoga Park, CA 91303 (U.S.A.)
Abstract
The potential for improved performance in advanced space propulsion systems through innovative application of advanced materials is discussed. The importance of a high strength-to-weight ratio, together with high temperature capability is examined, as also is the need for high thermal conductivity in actively cooled structures. Many advanced space propulsion systems are hydrogen fueled, making hydrogen compatibility a major requirement. The performance of candidate materials in high pressure hydrogen environments is discussed at length, and promising systems are identified.
I. Introduction
Materials for advanced space propulsion systems require properties similar to those used for conventional aircraft propulsion systems, i.e. high temperature capability and low density. In addition, however, in systems where hydrogen is used for fuel, hydrogen embrittlement resistance is also required. In advanced space propulsion systems currently being examined in the U.S.A., Europe and Japan, hydrogen is frequently considered as the prime candidate for the fuel, making hydrogen embrittlement resistance a necessary requirement for most of the materials considered. In addition, advanced space propulsion systems frequently require the materials to be cooled by the fuel during operation, making high thermal conductivity a useful attribute in designing materials for these systems. The reasons for these requirements will be dealt with in greater detail later in this paper. Over the last 30 years, significant advances have been made in improving the high temperature capabilities of materials designed for propulsion systems. Figure 1 shows the temperature for creep rupture in 1 0 0 h at 140 MN m -2 as a function of the year in which the materials were introduced. It can be seen that the temperature capability of propulsion materials over the last 35 years has advanced by more than 300 °C [1]. This improvement in maximum temperature capability
has resulted in enormous improvements in efficiency and, consequently, improved performance and reduced fuel consumption. These improvements in materials capability translate directly to improved performance in advanced propulsion systems for space applications. Higher temperatures enable the propulsion system to operate at higher efficiencies and have also allowed the weight of these propulsion systems to be reduced significantly. For example, the thrust-to-weight ratio of conventional gas turbine engines has been increased by a factor of 10 over the last 35 years, largely because of increases in alloy performance. Similar improvements in materials capability
1,200 O o...
A D V A N C E D SC
~ 1,100
~
1000
g ~. 2
900
~ ~o 17OO
M 246
~
194o
WASPALOY
B I
I
195o
19~o
197o
l~eo
'
YEAR
Fig. l. Temperature capability plotted in terms of temperature for 100 h rupture vs. year of introduction.
Elsevier Sequoia/Printed in T h e Netherlands
22
have enabled space propulsion systems to achieve high levels of thrust for given engine weight and improved fuel efficiency. For example, Fig. 2 shows two rocket engines having similar thrust levels with the larger engine on the left having a thrust-to-weight ratio of 20 while the smaller engine on the right has a thrust-to-weight ratio of 1200. Both engines have a thrust capabil-
ity of the order of 300 lbf. Furthermore, cost reductions can be achieved through innovative use of advanced materials and this is illustrated by the current example with the larger engine on the left costing around U.S.S150000 while the smaller engine on the right is estimated to cost around U.S.S3000 when each is produced in quantity [2]. It can be seen therefore that the innovative use of advanced materials can be a significant driver in rocket engine design performance and cost. Advanced space propulsion systems will be required for a new generation of hypersonic aircraft which are currently being considered. Hydrogen fuel is the propellant of choice because of its high efficiency, but this choice creates a wide range of materials challenges, including hydrogen embrittlement of structural materials and the need for high thermal conductivity where hydrogen is used to cool engine components. In addition, cryogenic properties must be adequate in order for engine components to survive the cool-down cycle prior to engine start. All these requirements are the focus of current studies and some of this work will be discussed in the following paragraphs.
2. Hydrogen embrittlement of advanced propulsion system materials
Fig. 2. Progress in propulsion technology illustrated by a comparison between conventional design on the left with a thrust-to-weight ratio of 20 and an advanced design on the right with a thrust-to-weight ratio of 1200 [2].
Hydrogen embrittlement may be generally characterized as being due to internal hydrogen embrittlement, to hydrogen reaction embrittlement or to hydrogen environment embrittlement as indicated in Table 1 [3]. Internal hydrogen embrittlement is due to hydrogen absorbed into
TABLE 1 Hydrogen embrittlement categories and characteristics Type of H embrittlement
Effects
Characteristics
Examples
Internal H embrittlement
H absorption at high temperatures and long exposure times Loss of strength and ductility
Worst at high temperatures
Almost all alloys except AI, Cu and Ag alloys
H reaction embrittlement
Hydride formation H~O vapor formation C~In formation
All alloys contain elements which react with H 2
Ti alloys C steels
H environment embrittlement
Loss of ductility Surface cracking Accelerated crack growth
Worst at approximately room temperature
Many Ni-base superalloys
23 the material, generally over a long period of time. Hydrogen reaction embrittlement is embrittlemerit due to hydrogen which reacts with elements in the material to form compounds such as a reaction with the carbon in steel to form methane. Hydrogen environment embrittlement is due to the effect of gaseous hydrogen at the material surface without significant absorption of hydrogen into the material but, nevertheless, causing a degradation in material properties. Depending on the material, hydrogen exposure conditions and mission duration, any or all of these three types of embrittlement have to be addressed in propulsion system design. For advanced propulsion systems where long-term service is frequently a requirement, internal hydrogen embrittlement can become of paramount importance whereas, in prior systems where only a single mission was required over a short period of time, this type of embrittlement was never a concern. Advanced space propulsion systems, on the contrary, must be capable of surviving hydrogen exposure for longer periods of time and frequently at higher temperatures. Table 2 shows a comparison of hydrogen fuel propulsion system characteristics for the A p o l l o J-2 system, the Space Shuttle main engine and future hypersonic propulsion systems. Durations of many hours at maximum hydrogen pressures and temperatures, which exceed those of any previous system, can be expected. Materials for hypersonic propulsion systems, examples of which are shown in Fig. 3, have the additional complication in that weight is extremely critical. The fuel fraction must be
increased to a maximum and structural weight to a minimum if mission requirements are to be achieved.
3. High conductivity composites A variety of materials will be required in order to satisfy these requirements, including high conductivity, copper-based composites, beryllium, aluminum, titanium alloys and superalloys. Figure 4 illustrates where some of these materials might be applied to such a vehicle. One of the critical applications is in high heat flux areas in the engine where high thermal conductivity is absolutely essential in order to reduce the thermal gradient through actively cooled compounds. Figure 5 shows the thermal conductivity of a range of high conductivity materials applicable to hypersonic propulsion systems. It can be seen that the thermal conductivity varies by a factor of at least 5. In terms of temperature gradient through the wall of an actively cooled component in a high heat flux environment, this difference
CONDITIONS: (DEPENDENT ON COOLANT ROUTING)
LOW HEAT FLUX MODERATE TEMP. BERYLLIUM TITANIUM ALUMINUM
MATERIALS:
HIGH HEAT FLUX LOW TEMP.
LOW HEAT FLUX HIGH TEMP,
C O P P E RALLOYS C O P P E RCOMPOSITES SUPERALLOYS BERYLLIUM
Fig. 3. Applications and characteristics of candidate materials for hypersonicpropulsion.
TABLE 2 Hydrogenfueledpropulsionsystemcharacteristics System
Application
Propel-Type lants
Apollo J-2
SaturnIB Saturn V
H 2, 0 2
Space Shuttle
Space Shuttle
U2, 0 2
main engine Hypersonic -propulsion
Dry Thrust SpecificDuration engine (at impulse(s) weight altitude) (s) (lbf) (lbf)
Gas generator 3454 230000 425 turbo pump-fed rocket engine Stagedcombustion 6335 470 000 455 turbo pump-fed rocket engine
H2, air --
500
Combustion Maximumseverity chamber of H exposure pressure (lbfin-z) PressureTemperature (lbfin-2) (°F) 780
1250
-~50
480 6000 (per FLT) (pre7.5 h total burners)
6000
750
9
9
--
9
24
- LEADING
EDGES
Fig. 4. Applicationsof high thermal conductivitymaterials for hypersonicpropulsion.
P100 50 Wo / Cu I0°1 K,, OXYGEN-FREE COPPER
2 m
COPPER MICROCOMPOSITE • LONGITUDINAL DIRECTION - - COOLING HEATING • SHORT TRANSVERSE COOLING / HEATING
200
a
~ 100
PIO0 50 r i o / Cu [0 °] Ke) P100 50 v/o / Cu [0°l K,~
0 -400
'
I 0
'
I 400
'
I 800
'
I 1200
1600
T E M P E R A T U R E , °F Fig. 5. T h e r m a l conductivities of various high conductivity materials.
translates into differences of several hundred degrees Fahrenheit through the wall of the heat exchanger and also into major changes in life of such a component. Figure 6 is the result of calculations of thermal gradient through a heat exchanger wall and several different materials and illustrates this effect, where the thermal gradient can vary from over 1000 K for a TiB2-TiA1 composite to less than 200 K for a high conductivity copper composite [4].
Materials which are candidates for such applications include beryllium, Cu-graphite composites, C u - N b microcomposites and conventional copper alloys. Evaluation of each of these systems in hydrogen is an important consideration in selecting the appropriate material. For example, the longitudinal ultimate strength of Cu-graphite composites is of significant interest because of the high strengths achieved over a wide range of temperatures, as shown ifl Fig. 7. The hydrogen performance of these com-
25
HEAT FLUX, W/m2X10 7 0
0.5
1.0
1.5
I
I
I
1200
Dimensions in inches
600 500 -
I-<1 I-
ujZ 400 -
o
1000
800
-4
£13 <
Y
F
300 < 200 -
-.4 I---
600
400 1
P130 X Gr(47)-Cu 100% INPLANE
TiB2(35)_TiA I
200 ~
100 -
/
Be
. . . . . . . . . . . . Cu-15%Nb THRU~ .................. 7)-Cu ° / °001
0 -!
0-
jICKNESS
0
500
1000
1500
HEAT FLUX, Btu/(fl 2 sec)
Fig. 6. Temperature gradient for materials having different thermal conductivities [4]•
"~
200
200
.c
¢=
"~ ~"
[]
As Received
[]
5 Cycles Inert
e..
[]
5 Cycles H2
100
-~
100
r4)
I.= 4) "~' o,
E
o -320
70 Temperature,
1000
1200
°F
Fig. 7. Longitudinal ultimate strengths of Cu-graphite composites (as-received material composite of 42 vol.% IM6 (P 100) carbon in an OFHC copper matrix, 0 ° laminate).
posites has been evaluated by cycling between room temperature and 1200 °F in high pressure hydrogen, and a low sum degradation in strength compared with materials cycled in an inert atmosphere has been observed. The general level of strength retention is excellent as shown in Fig. 8 and no methane formation has been observed. In general, copper alloys are very resistant to hydrogen embrittlement and, in fact, copper has been used as a coating to protect other materials from hydrogen exposure. Certain copper alloys can, however, display hydrogen reaction embrittlement due to the presence of sensitive alloy additions. For example, copper alloys containing excessive oxygen can be embrittled in high temperature hydrogen owing to internal water vapor
0
Fig. 8. Strength of Cu-graphite composite after cycling between room temperature and 1200 °F in inert gas and in hydrogen at a pressure of 3500 lbf in 2 (material composite of 42 vol.% IM6 (PAN) graphite fibers in an OFHC copper matrix, 0 ° laminate; straight-sided tensile specimens; edges coated with 0.003 in OFHC copper). The numbers of test specimens per condition are indicated in parentheses.
formation. In addition, certain microcomposites such as Cu-Nb can suffer from hydrogen embrittlement due to hydriding of the niobium.
4. Beryllium Beryllium has the unique advantage that it is light weight and virtually immune to hydrogen embrittlement. In addition, beryllium has a low hydrogen solubility and permeability as shown in Fig. 9. The permeability of hydrogen through beryllium is several orders of magnitude lower than through titanium, making beryllium a can-
26
o
II~TIITANIIuM
10 -2 E
10 . 4
~
10~
~
10-8
<
10-1o
m 0.
10-1:,
[_i
~",--~BERYLLIU , 7
8
.
,
,
,
9
10
11
1/T [l/K]
.
,
,
12
13
.
, 14
.
M
,~"~ 15
16
10000
Fig. 9. Hydrogen permeabilities of metals.
didate for coating other materials to protect them from internal hydrogen embrittlement. One problem, however, that will be difficult to address in the use of beryllium in cryogenic systems is its lack of low temperature ductility and fracture toughness. In order to satisfy a leak before burst requirement in pressurized systems, a minimum fracture toughness of the order of 20 MN m-2 will have to be achieved. Although modest improvements in fracture toughness by alloying are probably achievable, significant improvements will have to be recognized in order to make beryllium useful as a hypersonic propulsion system in advanced space propulsion systems.
5. Titanium alloys Although titanium alloys have been successfully applied in space structures for many years, the limitation that titanium has in terms of hydrogen resistance will make it difficult to find suitable applications in advanced space propulsion systems. Titanium has the advantage in that it has a low density and high strength combined with a good fracture toughness, and so a significant amount of effort has been expended to modify its properties to make it suitable for such applications. Titanium alloys have been used for a number of years for service in pumps and other components in high pressure cryogenic hydrogen. This is feasible because the low operating temperatures are such that the kinetics of hydrogen absorption and transport in the material are slow enough that embrittlement is not excessive. One always has to ensure, however, that the material is used in the relatively low stress applications where yielding is not a factor, and also where the
temperature is limited to room temperature or below. Recent data indicate that both Ti3Al-based and TiAl-based alloys are susceptible to hydrogen reaction embrittlement [5, 6]. It is still possible, however, that these intermetallic systems might be useful in high pressure hydrogen if the temperature is limited to modest levels. Alloy design approaches and coating systems may also be developed, but these alloys are not being considered for high temperature, high pressure hydrogen applications at the present time. Titanium alloys and intermetallic systems may be of use in casual hydrogen where partial pressures of hydrogen mixed with other gases are to be used. However, hydrogen absorption and hydrogen environment embrittlement are still a possibility and must be guarded against, using appropriate coatings.
6. Superalloys Iron-, nickel- and cobalt-based superalloys are all affected by hydrogen environment embrittlement to varying extents. Figure 10 shows the notched tensile strength ratio and smooth tensile ductility ratio for a variety of alloys. It can be seen that the extent of hydrogen embrittlement varies widely with alloy systems, but the degree of embrittlement roughly correlates with alloy strength. A number of alloys such as A286 and Incoloy 903 show almost no embrittlement. However, even these alloys may be affected by internal hydrogen embrittlement, resulting from long-term charging. Figure 11 shows the affect of hydrogen content on Incoloy 903 and Incoloy 718. Elongation and reduction in area after exposure times varying from 0 to 8 h at a high temperature are plotted, and it can be seen that significant degradations in elongation are noted even though the internal hydrogen levels are no greater than about 50 wt.ppm H 2. Superalloy structures operating at elevated temperatures in high pressure hydrogen can absorb significant amounts of hydrogen, and therefore the designer must be cognizant of the potential for reduced mechanical properties in such structures. Successful design requires attention to critical properties in the appropriate environment. Many years of satisfactory use in space propulsion components has proved that such items as turbine blades, high pressure fuel
27 1.2
-I-t • •
1.0 0.8
NOTCHEDTENSILESTRENGTHRATIO UNNOTCHEDTENSILEDUCTILITYRATIO (REDUCTIONOF AREA)
J •
1
-
-
1]
~
~ 0.4 0.2 0.0
~ ,"~"O"v " " × ' J" ~ " ~ ~b'~¢~ r"~ ~,'c C' - -~K '~~" 'o ,v~,"~"q, r°~ \~"."9'¢o x,,'L /,~",,.x,,q" ~'~"9'"< :..-k'" ~:~",:"9 ~ ,O ,%~,,,~ ~6,, c,,~"'"-', ,~
Fig. I 0. Notched tensile ratios at room temperature.
4(" "~
I He I H2
< 31 n-
o .~ E
c ~ 21
a.
Jq v
0
15 MIN Exposure Time
8 HR
0
15 MIN Exposure Time
8 HR
Fig. 11. Room temperature tensile properties of hydrogen-charged (1200 °F; 5000 lbf in 2; 15 Tin or 8 h) lncoloy 718 and Incoloy 903 tested at room temperature in H2-He at 5000 lbf in 2 (smooth tensile rounds; strain rate of 0.005 in in- ~ Tin m ( 8 . 3 3 × 1 0 5s I)tofailure).
lines and nozzles attest to the fact that such applications are entirely possible. However, in future systems, exposure times may exceed those in our current experience. It must be emphasized that a clear definition of the anticipated environment is a prerequEsite, together with careful testing and evaluation where the anticipated exposure exceeds current experience.
7. Ceramic composite systems Where higher temperatures are to be encountered that exceed the capabilities of current metallic systems, designers are beginning to consider alternatives such as ceramic matrix composites. Their high temperature capability makes ceramics and, in particular, ceramic composites and C - C composites a prime candidate for propulsion system applications. The deficiency that each of these systems has is a low fracture toughness and one of the approaches that has been taken recently is to
improve the level of fracture toughness of this type of material by introducing long fibers. The advantage of this approach is that fiber pull-out significantly increases fracture toughness and the resistance to crack propagation from small flaws. Analysis of this approach, showing the improveTent that can be generated, both analytically and experimentally, can be found in the literature [7]. The advantages of this approach in terms of temperature capability and strength-to-weight ratio are plotted in Fig. 12 and it can be seen that the peak temperature which can be achieved through this approach is significantly higher than that obtainable even with the best metal matrix composites. Candidate fibers for reinforcement of ceramic composites are numerous and some of the commercially available candidates are listed in Table 3. The goal properties which are desirable in a composite system are a coefficient of thermal expansion which closely matches the matrix, together with strength and modulus values which
28
satisfy the requirements. Generally speaking, high temperature applications require a stable fiber such as quartz or alumina. Fibers having multicomponent systems such as the complex A12O~-SiO 2 fibers generally do not retain their strength at elevated temperatures. The desirable fracture toughness of fiber-reinforced ceramic matrix composites results from sliding of the fibers within the matrix as the component is loaded and subcritical crack propagation occurs. A bond strength, which allows limited sliding and results in dissipation of energy when a crack propagates through the matrix, provides optimum
4000 (22051 u. o
3000 11650) D.
E
2000 o~ 11095) c
properties. Consequently, fibers are generally coated with an engineered interface to reduce chemical interaction between the reinforcement and the matrix and to encourage limited sliding. The micrograph shown in Fig. 13 is an example of the type of sliding that results in a typical fracture surface in such a composite system. For advanced propulsion system components such as rocket engine turbine blades, non-oxide ceramic matrixes combine good mechanical properties and good resistance to oxidation and high temperature. Both silicon carbide and silicon nitride have been used to fabricate ceramic matrix composites with various types of fiber, including carbon and silicon carbide. Hydrogen effects on ceramic matrix composites have not been investigated to any great extent, but there is a need to begin to study this area, both from a hydrogen permeability point of view, i.e. the hydrogen reaction, and also from the potential effects of hydrogen on the mechanical properties. Certain components of ceramic matrix composites such as silicon carbide can be reduced by hydrogen and elevated temperatures, making these systems potentially undesirable in such an application.
0.
o
1000 1540)
8. Summary 10 3
10 4 Strength/Weight Ratio (Inches or mm x 25.4)
Fig. 12. Strength-to-weight ratio of various high temperature materials plotted vs. temperature.
It can be seen from the foregoing discussion that a wide variety of materials will be used in advanced space propulsion systems. Generally speaking, since hydrogen is the propellant of choice, these materials must be resistant to
TABLE 3 Commercially available ceramic fibers Manufacturer
Trade name
Composition
Phases
DuPont DuPont
FP PRD- 166
AI20 s Al203-ZrO 2
Fiber Materials, Inc. ICI America, Inc. Owens Corning Textron Textron 3M Corporation 3M Corporation British Petroleum (U.K.) Nippon Carbon (Japan) UBE (Japan)
-Saffild E Glass SCS-6 CVD Boron Nextel 312 Nextel 480 Sigma Nicalon Tyranno
SiO~ A[203-SiO 2 SiO2-AIzO3-B203 SiC on C core B on C core AI,.O~-SiO:-B20~ A1203-SiO2-B:O~ SiC on W core SiC Si-Ti-C-O
a-A1203 a-Al203, tetragonal ZrO 2 Amorphous ¢~-A1203, mullite Amorphous SiC, C B, C 9A-2B, glass Mullite SiC, W SiC Amorphous
Diameter (/~m) 20 20 9-15 3 10 140 100, 140 8-12 10-12 100 10-20 8-10
Strength (MPa)
Modulus (GPa)
1380 2108
966 380
3400 2000 3450 3900 3585 1720 2285 3600 2900 2740
-300 72 415 422 152 225 420 195 206
29
Fig. 13. Fractory surface of a carbon-reinforced silicon carbide test specimen showing fiber pull-out.
hydrogen embrittlement and any reaction with hydrogen at elevated temperatures. In choosing a material for a given system, this then becomes a primary reconsideration and a first screening point for the choice of materials for such applications. Other property requirements such as fracture toughness, high temperature strength and creep properties then become of secondary importance in that hydrogen compatibility is a prerequisite and other properties must meet minimum requirements, once one has selected materials which are basically compatible with the appropriate temperatures and pressures of hydrogen to be anticipated in the system.
in providing some of the figures for this paper.
Acknowledgments
NASA-MSt:(" Advanced Earth-to-Orbit Propulsion Technolo,gy ('onJerence 1986, National Aeronautics and Space Administration-MSFC, Huntsville, AL, 1986. 7 A . G . Evans, Mater. Ak'i. Eng., A ( 1991 ).
The author is grateful to Dr. J. Yuen, Dr. L. Fritzemeier and Dr. D. Matejczyk, for assistance
References lnvestrnent-(klst Superalloys Advanced Materials and l'rocesses, August 199(), 1990, pp. 23 30. 2 W. Burns, Proof of principle: the kinetic kill vehicle, ThreshoM 6, 1990, pp. 2(I-29. 3 D. E. Matejczyk et al., The effect of high pressure hydrogen on materials for hypersonic propulsion applications, 8th Worhl ttydrogen Energy ('onJl, ttawaii, July 199(/, 1 W. J. Molloy,
199(1. 4 D. E. Matejczyk and C. G. Rhodes, Scr. Metall., 24 (7) (1990) 1369-1374. 5 D. S. Shih, I. M. Robertson and H. K. Birnbaum, Acta Metall.. 36 ( 1 ) (1988) 111 - 124. 6 W.T. Chandler, in R. J. Richmond and S. T. Yu (eds.), I'roc.
21
Materials for advanced space propulsion systems Neil E. Paton Rockwell International Corporation, Roeketdyne Division, 6633 Canoga Avenue, Canoga Park, CA 91303 (U.S.A.)
Abstract
The potential for improved performance in advanced space propulsion systems through innovative application of advanced materials is discussed. The importance of a high strength-to-weight ratio, together with high temperature capability is examined, as also is the need for high thermal conductivity in actively cooled structures. Many advanced space propulsion systems are hydrogen fueled, making hydrogen compatibility a major requirement. The performance of candidate materials in high pressure hydrogen environments is discussed at length, and promising systems are identified.
I. Introduction
Materials for advanced space propulsion systems require properties similar to those used for conventional aircraft propulsion systems, i.e. high temperature capability and low density. In addition, however, in systems where hydrogen is used for fuel, hydrogen embrittlement resistance is also required. In advanced space propulsion systems currently being examined in the U.S.A., Europe and Japan, hydrogen is frequently considered as the prime candidate for the fuel, making hydrogen embrittlement resistance a necessary requirement for most of the materials considered. In addition, advanced space propulsion systems frequently require the materials to be cooled by the fuel during operation, making high thermal conductivity a useful attribute in designing materials for these systems. The reasons for these requirements will be dealt with in greater detail later in this paper. Over the last 30 years, significant advances have been made in improving the high temperature capabilities of materials designed for propulsion systems. Figure 1 shows the temperature for creep rupture in 1 0 0 h at 140 MN m -2 as a function of the year in which the materials were introduced. It can be seen that the temperature capability of propulsion materials over the last 35 years has advanced by more than 300 °C [1]. This improvement in maximum temperature capability
has resulted in enormous improvements in efficiency and, consequently, improved performance and reduced fuel consumption. These improvements in materials capability translate directly to improved performance in advanced propulsion systems for space applications. Higher temperatures enable the propulsion system to operate at higher efficiencies and have also allowed the weight of these propulsion systems to be reduced significantly. For example, the thrust-to-weight ratio of conventional gas turbine engines has been increased by a factor of 10 over the last 35 years, largely because of increases in alloy performance. Similar improvements in materials capability
1,200 O o...
A D V A N C E D SC
~ 1,100
~
1000
g ~. 2
900
~ ~o 17OO
M 246
~
194o
WASPALOY
B I
I
195o
19~o
197o
l~eo
'
YEAR
Fig. l. Temperature capability plotted in terms of temperature for 100 h rupture vs. year of introduction.
Elsevier Sequoia/Printed in T h e Netherlands
22
have enabled space propulsion systems to achieve high levels of thrust for given engine weight and improved fuel efficiency. For example, Fig. 2 shows two rocket engines having similar thrust levels with the larger engine on the left having a thrust-to-weight ratio of 20 while the smaller engine on the right has a thrust-to-weight ratio of 1200. Both engines have a thrust capabil-
ity of the order of 300 lbf. Furthermore, cost reductions can be achieved through innovative use of advanced materials and this is illustrated by the current example with the larger engine on the left costing around U.S.S150000 while the smaller engine on the right is estimated to cost around U.S.S3000 when each is produced in quantity [2]. It can be seen therefore that the innovative use of advanced materials can be a significant driver in rocket engine design performance and cost. Advanced space propulsion systems will be required for a new generation of hypersonic aircraft which are currently being considered. Hydrogen fuel is the propellant of choice because of its high efficiency, but this choice creates a wide range of materials challenges, including hydrogen embrittlement of structural materials and the need for high thermal conductivity where hydrogen is used to cool engine components. In addition, cryogenic properties must be adequate in order for engine components to survive the cool-down cycle prior to engine start. All these requirements are the focus of current studies and some of this work will be discussed in the following paragraphs.
2. Hydrogen embrittlement of advanced propulsion system materials
Fig. 2. Progress in propulsion technology illustrated by a comparison between conventional design on the left with a thrust-to-weight ratio of 20 and an advanced design on the right with a thrust-to-weight ratio of 1200 [2].
Hydrogen embrittlement may be generally characterized as being due to internal hydrogen embrittlement, to hydrogen reaction embrittlement or to hydrogen environment embrittlement as indicated in Table 1 [3]. Internal hydrogen embrittlement is due to hydrogen absorbed into
TABLE 1 Hydrogen embrittlement categories and characteristics Type of H embrittlement
Effects
Characteristics
Examples
Internal H embrittlement
H absorption at high temperatures and long exposure times Loss of strength and ductility
Worst at high temperatures
Almost all alloys except AI, Cu and Ag alloys
H reaction embrittlement
Hydride formation H~O vapor formation C~In formation
All alloys contain elements which react with H 2
Ti alloys C steels
H environment embrittlement
Loss of ductility Surface cracking Accelerated crack growth
Worst at approximately room temperature
Many Ni-base superalloys
23 the material, generally over a long period of time. Hydrogen reaction embrittlement is embrittlemerit due to hydrogen which reacts with elements in the material to form compounds such as a reaction with the carbon in steel to form methane. Hydrogen environment embrittlement is due to the effect of gaseous hydrogen at the material surface without significant absorption of hydrogen into the material but, nevertheless, causing a degradation in material properties. Depending on the material, hydrogen exposure conditions and mission duration, any or all of these three types of embrittlement have to be addressed in propulsion system design. For advanced propulsion systems where long-term service is frequently a requirement, internal hydrogen embrittlement can become of paramount importance whereas, in prior systems where only a single mission was required over a short period of time, this type of embrittlement was never a concern. Advanced space propulsion systems, on the contrary, must be capable of surviving hydrogen exposure for longer periods of time and frequently at higher temperatures. Table 2 shows a comparison of hydrogen fuel propulsion system characteristics for the A p o l l o J-2 system, the Space Shuttle main engine and future hypersonic propulsion systems. Durations of many hours at maximum hydrogen pressures and temperatures, which exceed those of any previous system, can be expected. Materials for hypersonic propulsion systems, examples of which are shown in Fig. 3, have the additional complication in that weight is extremely critical. The fuel fraction must be
increased to a maximum and structural weight to a minimum if mission requirements are to be achieved.
3. High conductivity composites A variety of materials will be required in order to satisfy these requirements, including high conductivity, copper-based composites, beryllium, aluminum, titanium alloys and superalloys. Figure 4 illustrates where some of these materials might be applied to such a vehicle. One of the critical applications is in high heat flux areas in the engine where high thermal conductivity is absolutely essential in order to reduce the thermal gradient through actively cooled compounds. Figure 5 shows the thermal conductivity of a range of high conductivity materials applicable to hypersonic propulsion systems. It can be seen that the thermal conductivity varies by a factor of at least 5. In terms of temperature gradient through the wall of an actively cooled component in a high heat flux environment, this difference
CONDITIONS: (DEPENDENT ON COOLANT ROUTING)
LOW HEAT FLUX MODERATE TEMP. BERYLLIUM TITANIUM ALUMINUM
MATERIALS:
HIGH HEAT FLUX LOW TEMP.
LOW HEAT FLUX HIGH TEMP,
C O P P E RALLOYS C O P P E RCOMPOSITES SUPERALLOYS BERYLLIUM
Fig. 3. Applications and characteristics of candidate materials for hypersonicpropulsion.
TABLE 2 Hydrogenfueledpropulsionsystemcharacteristics System
Application
Propel-Type lants
Apollo J-2
SaturnIB Saturn V
H 2, 0 2
Space Shuttle
Space Shuttle
U2, 0 2
main engine Hypersonic -propulsion
Dry Thrust SpecificDuration engine (at impulse(s) weight altitude) (s) (lbf) (lbf)
Gas generator 3454 230000 425 turbo pump-fed rocket engine Stagedcombustion 6335 470 000 455 turbo pump-fed rocket engine
H2, air --
500
Combustion Maximumseverity chamber of H exposure pressure (lbfin-z) PressureTemperature (lbfin-2) (°F) 780
1250
-~50
480 6000 (per FLT) (pre7.5 h total burners)
6000
750
9
9
--
9
24
- LEADING
EDGES
Fig. 4. Applicationsof high thermal conductivitymaterials for hypersonicpropulsion.
P100 50 Wo / Cu I0°1 K,, OXYGEN-FREE COPPER
2 m
COPPER MICROCOMPOSITE • LONGITUDINAL DIRECTION - - COOLING HEATING • SHORT TRANSVERSE COOLING / HEATING
200
a
~ 100
PIO0 50 r i o / Cu [0 °] Ke) P100 50 v/o / Cu [0°l K,~
0 -400
'
I 0
'
I 400
'
I 800
'
I 1200
1600
T E M P E R A T U R E , °F Fig. 5. T h e r m a l conductivities of various high conductivity materials.
translates into differences of several hundred degrees Fahrenheit through the wall of the heat exchanger and also into major changes in life of such a component. Figure 6 is the result of calculations of thermal gradient through a heat exchanger wall and several different materials and illustrates this effect, where the thermal gradient can vary from over 1000 K for a TiB2-TiA1 composite to less than 200 K for a high conductivity copper composite [4].
Materials which are candidates for such applications include beryllium, Cu-graphite composites, C u - N b microcomposites and conventional copper alloys. Evaluation of each of these systems in hydrogen is an important consideration in selecting the appropriate material. For example, the longitudinal ultimate strength of Cu-graphite composites is of significant interest because of the high strengths achieved over a wide range of temperatures, as shown ifl Fig. 7. The hydrogen performance of these com-
25
HEAT FLUX, W/m2X10 7 0
0.5
1.0
1.5
I
I
I
1200
Dimensions in inches
600 500 -
I-<1 I-
ujZ 400 -
o
1000
800
-4
£13 <
Y
F
300 < 200 -
-.4 I---
600
400 1
P130 X Gr(47)-Cu 100% INPLANE
TiB2(35)_TiA I
200 ~
100 -
/
Be
. . . . . . . . . . . . Cu-15%Nb THRU~ .................. 7)-Cu ° / °001
0 -!
0-
jICKNESS
0
500
1000
1500
HEAT FLUX, Btu/(fl 2 sec)
Fig. 6. Temperature gradient for materials having different thermal conductivities [4]•
"~
200
200
.c
¢=
"~ ~"
[]
As Received
[]
5 Cycles Inert
e..
[]
5 Cycles H2
100
-~
100
r4)
I.= 4) "~' o,
E
o -320
70 Temperature,
1000
1200
°F
Fig. 7. Longitudinal ultimate strengths of Cu-graphite composites (as-received material composite of 42 vol.% IM6 (P 100) carbon in an OFHC copper matrix, 0 ° laminate).
posites has been evaluated by cycling between room temperature and 1200 °F in high pressure hydrogen, and a low sum degradation in strength compared with materials cycled in an inert atmosphere has been observed. The general level of strength retention is excellent as shown in Fig. 8 and no methane formation has been observed. In general, copper alloys are very resistant to hydrogen embrittlement and, in fact, copper has been used as a coating to protect other materials from hydrogen exposure. Certain copper alloys can, however, display hydrogen reaction embrittlement due to the presence of sensitive alloy additions. For example, copper alloys containing excessive oxygen can be embrittled in high temperature hydrogen owing to internal water vapor
0
Fig. 8. Strength of Cu-graphite composite after cycling between room temperature and 1200 °F in inert gas and in hydrogen at a pressure of 3500 lbf in 2 (material composite of 42 vol.% IM6 (PAN) graphite fibers in an OFHC copper matrix, 0 ° laminate; straight-sided tensile specimens; edges coated with 0.003 in OFHC copper). The numbers of test specimens per condition are indicated in parentheses.
formation. In addition, certain microcomposites such as Cu-Nb can suffer from hydrogen embrittlement due to hydriding of the niobium.
4. Beryllium Beryllium has the unique advantage that it is light weight and virtually immune to hydrogen embrittlement. In addition, beryllium has a low hydrogen solubility and permeability as shown in Fig. 9. The permeability of hydrogen through beryllium is several orders of magnitude lower than through titanium, making beryllium a can-
26
o
II~TIITANIIuM
10 -2 E
10 . 4
~
10~
~
10-8
<
10-1o
m 0.
10-1:,
[_i
~",--~BERYLLIU , 7
8
.
,
,
,
9
10
11
1/T [l/K]
.
,
,
12
13
.
, 14
.
M
,~"~ 15
16
10000
Fig. 9. Hydrogen permeabilities of metals.
didate for coating other materials to protect them from internal hydrogen embrittlement. One problem, however, that will be difficult to address in the use of beryllium in cryogenic systems is its lack of low temperature ductility and fracture toughness. In order to satisfy a leak before burst requirement in pressurized systems, a minimum fracture toughness of the order of 20 MN m-2 will have to be achieved. Although modest improvements in fracture toughness by alloying are probably achievable, significant improvements will have to be recognized in order to make beryllium useful as a hypersonic propulsion system in advanced space propulsion systems.
5. Titanium alloys Although titanium alloys have been successfully applied in space structures for many years, the limitation that titanium has in terms of hydrogen resistance will make it difficult to find suitable applications in advanced space propulsion systems. Titanium has the advantage in that it has a low density and high strength combined with a good fracture toughness, and so a significant amount of effort has been expended to modify its properties to make it suitable for such applications. Titanium alloys have been used for a number of years for service in pumps and other components in high pressure cryogenic hydrogen. This is feasible because the low operating temperatures are such that the kinetics of hydrogen absorption and transport in the material are slow enough that embrittlement is not excessive. One always has to ensure, however, that the material is used in the relatively low stress applications where yielding is not a factor, and also where the
temperature is limited to room temperature or below. Recent data indicate that both Ti3Al-based and TiAl-based alloys are susceptible to hydrogen reaction embrittlement [5, 6]. It is still possible, however, that these intermetallic systems might be useful in high pressure hydrogen if the temperature is limited to modest levels. Alloy design approaches and coating systems may also be developed, but these alloys are not being considered for high temperature, high pressure hydrogen applications at the present time. Titanium alloys and intermetallic systems may be of use in casual hydrogen where partial pressures of hydrogen mixed with other gases are to be used. However, hydrogen absorption and hydrogen environment embrittlement are still a possibility and must be guarded against, using appropriate coatings.
6. Superalloys Iron-, nickel- and cobalt-based superalloys are all affected by hydrogen environment embrittlement to varying extents. Figure 10 shows the notched tensile strength ratio and smooth tensile ductility ratio for a variety of alloys. It can be seen that the extent of hydrogen embrittlement varies widely with alloy systems, but the degree of embrittlement roughly correlates with alloy strength. A number of alloys such as A286 and Incoloy 903 show almost no embrittlement. However, even these alloys may be affected by internal hydrogen embrittlement, resulting from long-term charging. Figure 11 shows the affect of hydrogen content on Incoloy 903 and Incoloy 718. Elongation and reduction in area after exposure times varying from 0 to 8 h at a high temperature are plotted, and it can be seen that significant degradations in elongation are noted even though the internal hydrogen levels are no greater than about 50 wt.ppm H 2. Superalloy structures operating at elevated temperatures in high pressure hydrogen can absorb significant amounts of hydrogen, and therefore the designer must be cognizant of the potential for reduced mechanical properties in such structures. Successful design requires attention to critical properties in the appropriate environment. Many years of satisfactory use in space propulsion components has proved that such items as turbine blades, high pressure fuel
27 1.2
-I-t • •
1.0 0.8
NOTCHEDTENSILESTRENGTHRATIO UNNOTCHEDTENSILEDUCTILITYRATIO (REDUCTIONOF AREA)
J •
1
-
-
1]
~
~ 0.4 0.2 0.0
~ ,"~"O"v " " × ' J" ~ " ~ ~b'~¢~ r"~ ~,'c C' - -~K '~~" 'o ,v~,"~"q, r°~ \~"."9'¢o x,,'L /,~",,.x,,q" ~'~"9'"< :..-k'" ~:~",:"9 ~ ,O ,%~,,,~ ~6,, c,,~"'"-', ,~
Fig. I 0. Notched tensile ratios at room temperature.
4(" "~
I He I H2
< 31 n-
o .~ E
c ~ 21
a.
Jq v
0
15 MIN Exposure Time
8 HR
0
15 MIN Exposure Time
8 HR
Fig. 11. Room temperature tensile properties of hydrogen-charged (1200 °F; 5000 lbf in 2; 15 Tin or 8 h) lncoloy 718 and Incoloy 903 tested at room temperature in H2-He at 5000 lbf in 2 (smooth tensile rounds; strain rate of 0.005 in in- ~ Tin m ( 8 . 3 3 × 1 0 5s I)tofailure).
lines and nozzles attest to the fact that such applications are entirely possible. However, in future systems, exposure times may exceed those in our current experience. It must be emphasized that a clear definition of the anticipated environment is a prerequEsite, together with careful testing and evaluation where the anticipated exposure exceeds current experience.
7. Ceramic composite systems Where higher temperatures are to be encountered that exceed the capabilities of current metallic systems, designers are beginning to consider alternatives such as ceramic matrix composites. Their high temperature capability makes ceramics and, in particular, ceramic composites and C - C composites a prime candidate for propulsion system applications. The deficiency that each of these systems has is a low fracture toughness and one of the approaches that has been taken recently is to
improve the level of fracture toughness of this type of material by introducing long fibers. The advantage of this approach is that fiber pull-out significantly increases fracture toughness and the resistance to crack propagation from small flaws. Analysis of this approach, showing the improveTent that can be generated, both analytically and experimentally, can be found in the literature [7]. The advantages of this approach in terms of temperature capability and strength-to-weight ratio are plotted in Fig. 12 and it can be seen that the peak temperature which can be achieved through this approach is significantly higher than that obtainable even with the best metal matrix composites. Candidate fibers for reinforcement of ceramic composites are numerous and some of the commercially available candidates are listed in Table 3. The goal properties which are desirable in a composite system are a coefficient of thermal expansion which closely matches the matrix, together with strength and modulus values which
28
satisfy the requirements. Generally speaking, high temperature applications require a stable fiber such as quartz or alumina. Fibers having multicomponent systems such as the complex A12O~-SiO 2 fibers generally do not retain their strength at elevated temperatures. The desirable fracture toughness of fiber-reinforced ceramic matrix composites results from sliding of the fibers within the matrix as the component is loaded and subcritical crack propagation occurs. A bond strength, which allows limited sliding and results in dissipation of energy when a crack propagates through the matrix, provides optimum
4000 (22051 u. o
3000 11650) D.
E
2000 o~ 11095) c
properties. Consequently, fibers are generally coated with an engineered interface to reduce chemical interaction between the reinforcement and the matrix and to encourage limited sliding. The micrograph shown in Fig. 13 is an example of the type of sliding that results in a typical fracture surface in such a composite system. For advanced propulsion system components such as rocket engine turbine blades, non-oxide ceramic matrixes combine good mechanical properties and good resistance to oxidation and high temperature. Both silicon carbide and silicon nitride have been used to fabricate ceramic matrix composites with various types of fiber, including carbon and silicon carbide. Hydrogen effects on ceramic matrix composites have not been investigated to any great extent, but there is a need to begin to study this area, both from a hydrogen permeability point of view, i.e. the hydrogen reaction, and also from the potential effects of hydrogen on the mechanical properties. Certain components of ceramic matrix composites such as silicon carbide can be reduced by hydrogen and elevated temperatures, making these systems potentially undesirable in such an application.
0.
o
1000 1540)
8. Summary 10 3
10 4 Strength/Weight Ratio (Inches or mm x 25.4)
Fig. 12. Strength-to-weight ratio of various high temperature materials plotted vs. temperature.
It can be seen from the foregoing discussion that a wide variety of materials will be used in advanced space propulsion systems. Generally speaking, since hydrogen is the propellant of choice, these materials must be resistant to
TABLE 3 Commercially available ceramic fibers Manufacturer
Trade name
Composition
Phases
DuPont DuPont
FP PRD- 166
AI20 s Al203-ZrO 2
Fiber Materials, Inc. ICI America, Inc. Owens Corning Textron Textron 3M Corporation 3M Corporation British Petroleum (U.K.) Nippon Carbon (Japan) UBE (Japan)
-Saffild E Glass SCS-6 CVD Boron Nextel 312 Nextel 480 Sigma Nicalon Tyranno
SiO~ A[203-SiO 2 SiO2-AIzO3-B203 SiC on C core B on C core AI,.O~-SiO:-B20~ A1203-SiO2-B:O~ SiC on W core SiC Si-Ti-C-O
a-A1203 a-Al203, tetragonal ZrO 2 Amorphous ¢~-A1203, mullite Amorphous SiC, C B, C 9A-2B, glass Mullite SiC, W SiC Amorphous
Diameter (/~m) 20 20 9-15 3 10 140 100, 140 8-12 10-12 100 10-20 8-10
Strength (MPa)
Modulus (GPa)
1380 2108
966 380
3400 2000 3450 3900 3585 1720 2285 3600 2900 2740
-300 72 415 422 152 225 420 195 206
29
Fig. 13. Fractory surface of a carbon-reinforced silicon carbide test specimen showing fiber pull-out.
hydrogen embrittlement and any reaction with hydrogen at elevated temperatures. In choosing a material for a given system, this then becomes a primary reconsideration and a first screening point for the choice of materials for such applications. Other property requirements such as fracture toughness, high temperature strength and creep properties then become of secondary importance in that hydrogen compatibility is a prerequisite and other properties must meet minimum requirements, once one has selected materials which are basically compatible with the appropriate temperatures and pressures of hydrogen to be anticipated in the system.
in providing some of the figures for this paper.
Acknowledgments
NASA-MSt:(" Advanced Earth-to-Orbit Propulsion Technolo,gy ('onJerence 1986, National Aeronautics and Space Administration-MSFC, Huntsville, AL, 1986. 7 A . G . Evans, Mater. Ak'i. Eng., A ( 1991 ).
The author is grateful to Dr. J. Yuen, Dr. L. Fritzemeier and Dr. D. Matejczyk, for assistance
References lnvestrnent-(klst Superalloys Advanced Materials and l'rocesses, August 199(), 1990, pp. 23 30. 2 W. Burns, Proof of principle: the kinetic kill vehicle, ThreshoM 6, 1990, pp. 2(I-29. 3 D. E. Matejczyk et al., The effect of high pressure hydrogen on materials for hypersonic propulsion applications, 8th Worhl ttydrogen Energy ('onJl, ttawaii, July 199(/, 1 W. J. Molloy,
199(1. 4 D. E. Matejczyk and C. G. Rhodes, Scr. Metall., 24 (7) (1990) 1369-1374. 5 D. S. Shih, I. M. Robertson and H. K. Birnbaum, Acta Metall.. 36 ( 1 ) (1988) 111 - 124. 6 W.T. Chandler, in R. J. Richmond and S. T. Yu (eds.), I'roc.