![Soyuz rocket. Supplies of equipment and food are trans-
ported to the ISS using either the Russian Progress space-
craft of the European Automatic Transfer Vehicle. Since its
first flight in 2008, the ATV has played a vital role in ISS
logistics serving as a cargo carrier, ‘space tug’ and storage
facility. ATV will evolve into the European Service Module
(ESV) designed to support the NASA Orion spacecraft. The
deep-space missions envisaged by NASA will rely on the
Orion spacecraft—it has been flown on a successful test
flight aboard a Delta 4 Heavy booster launched from the
Kennedy Space Centre at the end of 2014.
Propellants for launch vehicles are regarded as materials.
All rockets are propelled into the vacuum of space by
utilizing only the liquid or solid materials on-board; at pre-
sent, there is no possibility of using atmospheric oxygen.
The term ‘propellant’ is used to denote the two chemical
products, the ‘oxidizer’ and the ‘fuel’, contained inside
conventional rockets. The fuel is burnt with the oxidizer in
order to achieve the enormous amounts of energy needed for
liftoff. The propellants can be in either liquid or solid state.
Modern rockets are powered by ‘cryogenic’ propellants:
liquids which at atmospheric pressure have boiling points
below 0 °C. Examples are liquid oxygen (−183 °C) and
liquid hydrogen (−253 °C).
The most simple solid propellant is prepared from a
mixture of nitrocellulose and nitroglycerine—this is the fuel.
Table 2.6 Selected launch vehicles
Country of origin Launch vehicle (latest known version) Typea
Payload into GTOb
(kg)
China Long March 3 (1992) ELV 2500
Long March 3B (2007) ELV 11,500 (LEO)
5500
European Space Agency Ariane 4 (1990) ELV 2600
Ariane 5 (1996) ELV 6800
Ariane 5 (2014) ELV 10,500
Vega—Italy (2014) ELV 1500 (LEO)
India Vehicle 3 (1979) ELV 40 (LEO)
Polar Satellite Launch Vehicle (2014) ELV 3250 (LEO)
Israel Shavit (1988) ELV 160 (LEO)
Japan H-1 (1986) ELV 1100
H-IIB (2009) ELV 16,500 (LEO)
8000
USA Scout (1979) ELV 5400
Atlas 2 (1991) ELV 2700
Atlas 531 (2014) ELV 17,000 (LEO)
Thor Delta (1992) ELV 2000
Delta 2 (2012) ELV 2500
Saturn V ELV 10,000
Titan III and IV (1989) ELV 5000
Falcon 9 (2014) ELV 13,150 (LEO)
Space Shuttle (1990, retired 2011) AV 25,000 (LEO)
Former USSR Vostok (1960) ELV 5000 (LEO)
Proton (1968) Russia ELV 5500
Soyuz-2.1b (2014) Russia ELV 3000
Soyuz-2.1b (2014) Russia ELV 8500 (LEO)
Soyuz-2.1v (2013) Russia ELV 2800
Zenit 3SL (2002) Ukraine ELV 6000
Buran Retired AV 30,000 (LEO)
Key
a
ELV expendable launch vehicle; AV aerospace vehicle
b
Geosynchronous transfer orbit [flight performance to Low Earth Orbit (LEO) is usually more than twice this payload weight]
2.4 Materials for Space Launch Vehicles 29](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-51-320.jpg)
![The oxidizer is prepared separately from either ammonium
nitrate, ammonium perchlorate, or potassium perchlorate.
The fuel and oxidizer are made into powder form and then
mixed with binder such as polyvinyl chloride or a poly-
urethane. The resulting substance is poured into the solid
rocket motor casing, where it sets hard. The case is usually
made of steel, as discussed previously for the Thor Delta
solid rocket motor, but composite materials based on gra-
phite fibres are also used. More complex solid propellant
chemistry is based on polybutaidene acrylonitrile (PBAN)
and hydroxytelechelique polybutadiene (HTPB) propellant.
The HTPB propellant for Ariane V solid boosters is, typi-
cally, 68 % ammonium perchlorate, 14 % polybutadiene,
and 18 % aluminium powder. Much progress has been made
during the last decade in the development of improved solid
propellants. One propellant is identified as GAP/Al/HNF
[glycidyl azide polymer being the binder, aluminium powder
and a new form of powerful oxidizer with the chemical
composition of hydrazinium nitroformate (Schoyer 1996)].
Comparative tests were made between this new propellant
and the best-performing HTPB-containing propellant. The
findings included an increase in characteristic velocity of
about 8 % and, importantly, an ecologically benign exhaust,
free of chlorine, where the combustion products are nitrogen,
water, carbon dioxide, nitrogen oxide, and aluminium oxide.
Liquid propellant motors are usually fed from two tanks,
one containing the liquid fuel (such as kerosene, liquid
hydrogen, or hydrazine (N2H4)), the other containing the
oxidizer (usually liquid oxygen or nitrogen tetroxide (N2O4)).
Infrequently, fuming red nitric acid or nitrogen peroxide is
used. The propellants are injected into the rocket motor’s
combustion chamber, where ignition and combustion occur
with a great release of thermal energy. The combustion gases
are then forced through the exit nozzle of the motor. Here, the
kinetic energy is absorbed by the nozzle as the gas velocity
increases and then decreases through the ‘throat’ of the
nozzle. A drawing of the Ariane IV launcher vehicle parts is
shown in Fig. 2.7. The first two stages of this ELV use
hydrazine and nitrogen tetroxide. The third stage is propelled
by a cryogenic engine fed by liquid hydrogen and liquid
oxygen. As illustrated in the figure, Ariane IV has the pos-
sibility to incorporate either liquid or solid propellant booster
motors to assist in the first stage liftoff.
In contrast to Ariane, the Space Shuttle was, as its name
implied, reusable and of extreme importance for its role in
manned, near-Earth activities. The Shuttle’s primary
propulsion consisted of a large external tank (47 m in length
and 8.7 m in diameter) which contained compartments for the
liquid oxygen/liquid hydrogen propellants—these were fed
to the three main engines, and two strap-on booster motors
which contained a solid composite propellant (polybutadiene,
acrylic binder, and ammonium perchlorate oxidizer).
Rocket structural materials are usually based on the
Duralumin series of aluminium alloys (4 %Cu, 2 %Mn, rem.
Al). These alloys have high strength-to-weight ratios and are
detailed in further sections of this book. Concerning Ari-
ane IV, much use has been made of the aluminium alloys
AU4GN (AA2024), AZ5GU (AA7075), and AZ5G
(AA7020). These are the French alloy designations with the
US Aluminium Association designation in parentheses.
Further cross-references to national alloy specifications are
given in Appendix 6. Some locations for these alloys are
indicated on Fig. 2.7. Unfortunately, several of their
heat-treatment conditions are susceptible to stress corrosion
cracking (SCC). A particular problem resulting from SCC is
discussed in Sect. 4.5. All the major structural materials used
for the construction of Ariane and its motors are identified in
Fig. 2.7. The more modern rocket motor and booster bodies
containing solid propellants are machined and welded from
‘maraging’ steels, which are based on iron with large
amounts of nickel, cobalt, and molybdenum. These steels
Fig. 2.6 View of the main structural parts of an ELV (based on a
Titan III design)
30 2 Requirements for Spacecraft Materials](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-52-320.jpg)
![been described as ‘highly automated’ (Isakowitz 1995).
General views of the Zenit assembly hall and its kerosene–
LOX engine are seen in Figs. 2.9 and 2.10 respectively.
This ELV has recently been proposed for launching satellites
from the sea. The novel idea is to utilize a covered
semi-submersible oil rig as the firing platform. This would
be anchored in the Pacific, at the Equator, so making full
advantage of the Earth’s maximum rotational velocity (about
1600 kph). The Zenit rocket, standing 62 m tall, would face
eastward in a direction avoiding any inhabitable landfall. All
other major world space launch sites are rather far north
from the Equator with the exception of ESA’s facility at
Kourou, French Guiana, which is less than 500 km north of
the Equator. The uniquely flexible ‘Sea Launch’ site will
undoubtedly require the implementation of a comprehensive
corrosion protection scheme for all the associated spacecraft
materials.
The Ariane V development was initiated by the European
Space Agency in 1985. It is designed for commercial mis-
sions to launch satellites and cargo for the future space
stations. The proposal for launch of the European Hermes as
a reusable winged manned space vehicle atop the Ariane V
was cancelled in the mid 90’s. However, it has flown several
cargo vehicle [Automated Transfer Vehicle (ATV)] and is
designed for a Crew Transfer Vehicle (CTV). An illustration
of the Ariane V in a dual launch configuration is shown in
Fig. 2.11. Ariane V has a length of 54 m, a gross mass of
710,000 kg, and a designed thrust at liftoff of 15.9 MN. The
Vulcain engine powers the Ariane V main cryogenic stage.
This engine consists of a gas generator cycle, in which tur-
bopumps driven by a gas generator fed by propellants tapped
from the main supply system feed fuel and oxidizer to the
combustion chamber. Liquid oxygen (oxidizer) and liquid
hydrogen (fuel) are sprayed into the combustion chamber.
Because of the extremely high combustion temperature,
reaching 3600 °C and with about 1600 °C at the internal wall
of the chamber, it is necessary to cool the chamber. This is
done by machining channels into the chamber wall—there
are 360—and passing liquid hydrogen through them as the
engine is fired. The chamber is fabricated from a wrought
high-strength copper alloy (Narloy-Z) with an outer band of
nickel. The combustion chambers of the Space Shuttle main
engines were also made from Narloy-Z (Cu, 3Ag, 0.5Zr wt%
with O2 of approximately 50 ppm) and, when correctly heat
treated, this alloy was ideal for combustion chambers oper-
ating from −252 to 540 °C but above this temperature the
hot wall mechanical properties were degraded because of
grain growth, grain boundary sliding and bulging that
Fig. 2.8 A liquid propellant rocket engine
Fig. 2.10 Installation of Zenit first stage kerosene—LO2 engine
Fig. 2.9 General view of the Zenit assembly hall
32 2 Requirements for Spacecraft Materials](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-54-320.jpg)
![principium individuationis, like a kaleidoscope, shows us in ever-
shifting evanescent forms, there is an underlying unity, not only truly
existing, but actually accessible to us; for lo! in tangible, objective
form, it stands before our sight.
Of these two mental attitudes, according as the one or the other is
adopted, so the ϕιλία (Love) or the νεῑκος (Hatred) of Empedocles
appears between man and man. If any one, who is animated by
νεῑκος, could forcibly break in upon his most detested foe, and
compel him to lay bare the inmost recesses of his heart; to his
surprise, he would find again in the latter his very self. For just as in
dreams, all the persons that appear to us are but the masked
images of ourselves; so in the dream of our waking life, it is our own
being which looks on us from out our neighbours' eyes,—though this
is not equally easy to discern. Nevertheless, tat tvam asi.
The preponderance of either mode of viewing life not only
determines single acts; it shapes a man's whole nature and
temperament. Hence the radical difference of mental habit between
the good character and the bad. The latter feels everywhere that a
thick wall of partition hedges him off from all others. For him the
world is an absolute non-ego, and his relation to it an essentially
hostile one; consequently, the key-note of his disposition is hatred,
suspicion, envy, and pleasure in seeing distress. The good character,
on the other hand, lives in an external world homogeneous with his
own being; the rest of mankind is not in his eyes a non-ego; he
thinks of it rather as myself once more. He therefore stands on an
essentially amicable footing with every one: he is conscious of being,
in his inmost nature, akin to the whole human race,[9]
takes direct
interest in their weal and woe, and confidently assumes in their case
the same interest in him. This is the source of his deep inward
peace, and of that happy, calm, contented manner, which goes out
on those around him, and is as the presence of a good diffused.
Whereas the bad character in time of trouble has no trust in the help
of his fellow-creatures. If he invokes aid, he does so without
confidence: obtained, he feels no real gratitude for it; because he
can hardly discern therein anything but the effect of others' folly. For](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-56-320.jpg)
![he is simply incapable of recognising his own self in some one else;
and this, even after it has furnished the most incontestible signs of
existence in that other person: on which fact the repulsive nature of
all unthankfulness in reality depends. The moral isolation, which thus
naturally and inevitably encompasses the bad man, is often the
cause of his becoming the victim of despair. The good man, on the
contrary, will appeal to his neighbours for assistance, with an
assurance equal to the consciousness he has of being ready himself
to help them. As I have said: to the one type, humanity is a non-
ego; to the other, myself once more. The magnanimous character,
who forgives his enemy, and returns good for evil, rises to the
sublime, and receives the highest meed of praise; because he
recognises his real self even there where it is most conspicuously
disowned.
Every purely beneficent act all help entirely and genuinely unselfish,
being, as such, exclusively inspired by another's distress, is, in fact,
if we probe the matter to the bottom, a dark enigma, a piece of
mysticism put into practice; inasmuch as it springs out of, and finds
its only true explanation in, the same higher knowledge that
constitutes the essence of whatever is mystical.
For how, otherwise than metaphysically, are we to account for even
the smallest offering of alms made with absolutely no other object
than that of lessening the want which afflicts a fellow-creature? Such
an act is only conceivable, only possible, in so far as the giver
knows that it is his very self which stands before him, clad in the
garments of suffering; in other words, so far as he recognises the
essential part of his own being, under a form not his own.[10]
It now
becomes apparent, why in the foregoing part I have called
Compassion the great mystery of Ethics.
He, who goes to meet death for his fatherland, has freed himself
from the illusion which limits a man's existence to his own person.
Such a one has broken the fetters of the principium individuationis.
In his widened, enlightened nature he embraces all his countrymen,
and in them lives on and on. Nay, he reaches forward to, and](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-57-320.jpg)
![merges himself in the generations yet unborn, for whom he works;
and he regards death as a wink of the eyelids, so momentary that it
does not interrupt the sight.
We may here sum up the characteristics of the two human types
above indicated. To the Egoist all other people are uniformly and
intrinsically strangers. In point of fact, he considers nothing to be
truly real, except his own person, and regards the rest of mankind
practically as troops of phantoms, to whom he assigns merely a
relative existence, so far as they may be instruments to serve, or
barriers to obstruct, his purposes; the result being an immeasurable
difference, a vast gulf between his ego on the one side, and the
non-ego on the other. In a word, he lives exclusively centred in his
own individuality, and on his death-day he sees all reality, indeed the
whole world, coming to an end along with himself.[11]
Whereas the
Altruist discerns in all other persons, nay, in every living thing, his
own entity, and feels therefore that his being is commingled, is
identical with the being of whatever is alive. By death he loses only a
small part of himself. Patting off the narrow limitations of the
individual, he passes into the larger life of all mankind, in whom he
always recognised, and, recognising, loved, his very self; and the
illusion of Time and Space, which separated his consciousness from
that of others, vanishes. These two opposite modes of viewing the
world are probably the chief, though not indeed the sole cause of
the difference we find between very good and exceptionally bad
men, as to the manner in which they meet their last hour.
In all ages Truth, poor thing, has been put to shame for being
paradoxical; and yet it is not her fault. She cannot assume the form
of Error seated on his throne of world-wide sovereignty. So then,
with a sigh, she looks up to her tutelary god, Time, who nods
assurance to her of future victory and glory, but whose wings beat
the air so slowly with their mighty strokes, that the individual
perishes or ever the day of triumph be come. Hence I, too, am
perfectly aware of the paradox which this metaphysical explanation
of the ultimate ethical phaenomenon must present to Western
minds, accustomed, as they are, to very different methods of](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-58-320.jpg)
![providing Morals with a basis. Nevertheless, I cannot offer violence
to the truth. All that is possible for me to do, out of consideration for
European blindness, is to assert once more, and demonstrate by
actual quotation, that the Metaphysics of Ethics, which I have here
suggested, was thousands of years ago the fundamental principle of
Indian wisdom. And to this wisdom I point back, as Copernicus did
to the Pythagorean cosmic system, which was suppressed by
Aristotle and Ptolemaeus. In the Bhagavadgîtâ (Lectio XIII.; 27, 28),
according to A. W. von Schlegel's translation, we find the following
passage: Eundem in omnibus animantibus consistentem summum
dominum, istis pereuntibus kaud pereuntem qui cernit, is vere cernit.
Eundem vero cernens ubique praesentem dominum, non violat
semet ipsum sua ipsius culpa: exinde pergit ad summum iter.[12]
With these hints towards the elaboration of a metaphysical basis for
Ethics I must close, although an important step still remains to be
taken. The latter would presuppose a further advance in Moral
Science itself; and this can hardly be made, because in the West the
highest aim of Ethics is reached in the theory of justice and virtue.
What lies beyond is unknown, or at any rate ignored. The omission,
therefore, is unavoidable; and the reader need feel no surprise, if
the above slight outline of the Metaphysics of Ethics does not bring
into view—even remotely—the corner-stone of the whole
metaphysical edifice, nor reveal the connection of all the parts
composing the Divina Commedia. Such a presentment, moreover, is
involved neither in the question set, nor in my own plan. A man
cannot say everything in one day, and should not answer more than
he is asked.
He who tries to promote human knowledge and insight is destined to
always encounter the opposition of his age, which is like the dead
weight of some mass that has to be dragged along: there on the
ground it lies, a huge inert deformity, defying all efforts to quicken
its shape with new life. But such a one must take comfort from the
certainty that, although prejudices beset his path, yet the truth is
with him. And Truth does but wait for her ally, Time, to join her;](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-59-320.jpg)
![once he is at her side, she is perfectly sure of victory, which, if to-
day delayed, will be won to-morrow.
[1] The conception of the Good, in its purity, is an ultimate one, an absolute Idea,
whose substance loses itself in infinity.—(Bouterweek: Praktische Aphorismen, p.
54.)
It is obvious that this writer would like to transform the familiar, nay, trivial
conception Good into a sort of Διἴπετής, to be set up as an idol in his temple.
Διἴπετής lit., fallen from Zeus; and so heaven-sent, a thing of divine origin.
Cf. Horn., Il.. XVI, 174; Od.. IV. 477. Eur., Bacch., 1268.—(Translator.)
[2]The genuineness of the O u p n e k ' h a t has been disputed on the ground of
certain marginal glosses which were added by Mohammedan copyists, and then
interpolated in the text, it has, however, been fully established by the Sanskrit
scholar, F.H.H. Windischmann (junior) in his Sancara, sive de Theologumenis
Vedanticorum, 1833, p. xix; and also by Bochinger in his book De la Vie
Contemplative chez les Indous, 1831, p. 12. The reader though ignorant of
Sanskrit, may yet convince himself that Anquetil Duperron's word for word Latin
translation of the Persian version of the U p a n i s h a d s made by the martyr of
this creed, the Sultan D â r â - S h u k o h, is based on a thorough and exact
knowledge of the language. He has only to compare it with recent translations of
some of the U p a n i s h a d s by Rammohun Boy, by Poley, and especially with that
of Colebrooke, as also with Röer's, (the latest). These writers are obviously
groping in obscurity, and driven to make shift with hazy conjectures, so that
without doubt their work is much less accurate. More will be found on this subject
in Vol. II. of the Parerga, chap. 16, § 184. [V. The Upanishads, translated by Max
Müller, in The Sacred Books of the East, Vols. I. and XV. Cf. also Max Müller, The
Science of Language, Vol. I., p. 171. Now that an adequate translation of the
original exists, the O u p n e k ' h a t has only an historical interest. The value which
Schopenhauer attached to the U p a n i s h a d s is very clearly expressed also in the
Welt als Wille und Vorstellung, Preface to the first Edition; and in the Parerga, II.,
chap, xvi., § 184.—(Translator.)]
[3] For the S û f i, more correctly *Sūfīy a sect which appeared already in the first
century of the H i j r a h, the reader is referred to: Tholuck's Blüthensammlung aus
der Morgenländischen Mystik (Berlin, 1825); Tholuck's Sûfismus, sive Theosophia
Persarum Pantheistica (Berlin, 1821); Kremer's Geschichte der Herrschenden
Ideen des Islâms (Leipzig, 1868); Palmer's Oriental Mysticism (London, 1867);
Gobineau's Les Religions et les Philosophies dans l'Asie Centrale (2nd edit. Paris,
1866); A Dictionary of Islâm, by T. P. Hughes (London, 1885), p. 608 sqq.—
(Translator.)](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-60-320.jpg)
![[4] This is too well-known to need verification by references. The Cantico del Sole
by St. Francis of Assisi sounds almost like a passage from the U p a n i s h a d s or
the B h a g a v a d g î t â.—(Translator.)
[5]
On peut assez longtemps, chez notre espèce,
Fermer la porte à la Raison.
Mais, dès qu'elle entre avec adresse,
Elle reste dans la maison,
Et bientôt elle en est maîtresse.
—(Voltaire.)
(We men may, doubtless, all our lives
To Reason bar the door.
But if to enter she contrives,
The house she leaves no more,
And soon as mistress there presides.)
[6] Τὸ ἔν= the eternal Reality outside Time and Space Tὸ πᾱν = the phaenomenal
universe.—(Translator.)
[7] Mâyâ is the delusive reflection of the true eternal Entity.—(Translator.)
[8] This expression is used in the Brahmanical philosophy to denote the relation
between the world-fiction as a whole and its individualised parts. V. A. E. Gough,
Philosophy of the Upanishads, 1882.—(Translator.)
[9] Homo sum: humani nil a me alienum puto. Terence, Heaut., I. 1, 25.—
(Translator.)
[10] It is probable that many, perhaps, most cases of truly disinterested
Compassion—when they really occur—are due not to any conscious knowledge of
this sort, but to an unconscious impulse springing from the ultimate unity of all
living things, and acting, so to say, automatically.—(Translator.)
[11] Cf. Richard Wagner: Jesus von Nazareth; pp. 79-90.—(Translator.)
[12] That man is endowed with true insight who sees that the same ruling power
is inherent in all things, and that when these perish, it perishes not. For if he
discerns the same ruling power everywhere present, he does not degrade himself
by his own fault: thence he passes to the highest path.—For the Bhagavadgîtâ the
reader is referred to Vol. VIII. of The Sacred Books of the East (Oxford: Clarendon
Press), where (p. 105) this passage is translated as follows:—He sees (truly) who
sees the supreme lord abiding alike in all entities, and not destroyed though they
are destroyed. For he who sees the lord abiding everywhere alike, does not
destroy himself[*] by himself, and then reaches the highest goal.
[*]Not to have true knowledge, is equivalent to self-destruction.](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-61-320.jpg)
![JUDGMENT OF THE DANISH ROYAL SOCIETY OF SCIENCES.
In 1837 the following question was set as subject for a Prize Essay:
Is the fountain and basis of Morals to be sought for in an idea of
morality which lies directly in the consciousness (or conscience), and
in the analysis of the other leading ethical conceptions which arise
from it? Or is it to be found in some other source of knowledge?
There was only one competitor; but his dissertation, written in
German, and bearing the motto: To preach Morality is easy, to
found it is difficult[1]
we cannot adjudge worthy of the Prize. He has
omitted to deal with the essential part of the question, apparently
thinking that he was asked to establish some fundamental principle
of Ethics. Consequently, that part of the treatise, which explains how
the moral basis he proposes is related to his system of metaphysics,
we find relegated to an appendix, as an opus supererogationis,
although it was precisely the connection between Metaphysics and
Ethics that our question required to be put in the first and foremost
place. The writer attempts to show that compassion is the ultimate
source of morality; but neither does his mode of discussion appear
satisfactory to us, nor has he, in point of fact, succeeded in proving
that such a foundation is adequate. Indeed he himself is obliged to
admit that it is not.[2]
Lastly, the Society cannot pass over in silence
the fact that he mentions several recent philosophers of the highest
standing in an unseemly manner, such as to justly occasion serions
offence.
[1] The Academy has been good enough to insert the second is on its own
account, by way of proving the truth of Longinus' theory (V. De Sublimitate: chap.
39, ad fin.), that the addition or subtraction of a single syllable is sufficient to
destroy the whole force of a sentence. (P. Longinus: De Sublimitate Libellus; edit.
Joannes Vablen, Bonnae, 1887.)—(Translator)
[2] I suppose this is the meaning of contra esse confiteri.— (Translator.)](https://image.slidesharecdn.com/2677135-250530074036-b0038593/85/Materials-And-Processes-For-Spacecraft-And-High-Reliability-Applications-1st-Edition-Barrie-D-Dunn-Auth-63-320.jpg)
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- 5. Materials and Processes for Spacecraft and High Reliability Applications Barrie D. Dunn
- 6. Springer Praxis Books Astronautical Engineering
- 7. More information about this series at http://www.springer.com/series/5495
- 8. Barrie D. Dunn Materials and Processes for Spacecraft and High Reliability Applications 123
- 9. Barrie D. Dunn School of Engineering University of Portsmouth Portsmouth UK Published in association with Praxis Publishing, Chichester, UK ISSN 2365-9599 ISSN 2365-9602 (electronic) Springer Praxis Books ISBN 978-3-319-23361-1 ISBN 978-3-319-23362-8 (eBook) DOI 10.1007/978-3-319-23362-8 Library of Congress Control Number: 2015948763 Springer Cham Heidelberg New York Dordrecht London © Springer International Publishing Switzerland 2016 This work is subject to copyright. All rights are reserved by the Publisher, whether the whole or part of the material is concerned, specifically the rights of translation, reprinting, reuse of illustrations, recitation, broadcasting, reproduction on microfilms or in any other physical way, and transmission or information storage and retrieval, electronic adaptation, computer software, or by similar or dissimilar methodology now known or hereafter developed. The use of general descriptive names, registered names, trademarks, service marks, etc. in this publication does not imply, even in the absence of a specific statement, that such names are exempt from the relevant protective laws and regulations and therefore free for general use. The publisher, the authors and the editors are safe to assume that the advice and information in this book are believed to be true and accurate at the date of publication. Neither the publisher nor the authors or the editors give a warranty, express or implied, with respect to the material contained herein or for any errors or omissions that may have been made. Cover design: Jim Wilkie Cover images: Front cover top—The Falcon 9 rocket streaks towards space from Florida’s Cape Canaveral Air Force Station containing supplies, including the first 3D printer in space and a troop of 20 mice, for the International Space Station (Courtesy SpaceX). Front cover lower—the assembly and integration of a satellite in SSTL’s clean-room (Courtesy of Surrey Satellite Technology Ltd.). Rear cover—Vega VV05 in its mobile gantry prior to launch at Europe’s Spaceport in Kourou, French Guiana (Courtesy ESA-M. Pedoussaut). Printed on acid-free paper Springer International Publishing AG Switzerland is part of Springer Science+Business Media (www.springer.com)
- 10. Talking of education, ‘People have now a-days, (said he,) got a strange opinion that everything should be taught by lectures. Now, I cannot see that lectures can do so much good as reading the books from which the lectures are taken. I know nothing that can be best taught by lectures, except where experiments are to be shewn. You may teach chemistry by lectures—You might teach making of shoes by lectures!’ Samuel Johnson, 1766 (from Boswell’s Life)
- 11. This book is dedicated to Cato and Dennis
- 12. Preface This book, as implied by the title page, is an extensively revised version of the former “Metallurgical Assessment of Spacecraft Parts, Materials and Processes” published in 1997. The present title has been modified to set it apart from the previous work and describe its expanded content. The book has become more voluminous, this reflects the huge advances made during the past 20 years when we have witnessed the increased usage of modern materials and manufacturing techniques that were unforeseeable when the former book was written. Also, the number of case studies and amount of general information has been extended to become a source for engineers, space scientists, laboratory experimenters and technicians. Although much of the book considers metallurgical aspects of spacecraft engi- neering, there is now basic advice covering organic and ceramic materials as well as tech- niques available for assembling them into essential sub-systems, reliable parts and structures. A good number of the original illustrations are retained but many new ones have been added. Several images reflect the quite remarkable outcomes of space projects. These include high resolution images of Earth taken by satellites which are relevant for surveillance and the forecasting of weather. Also included are fly-by images of enigmatic little moons and comets captured by spacecraft after many years of voyaging in search of life and the origins of water in our own Solar System. Equipment on-board the International Space Station and satellite-based communications are mentioned. These have all been made possible by breakthroughs in materials, processes and electronic-engineering. Plato saw engineers as “doers” not “thinkers”. From ancient times no one expected engi- neers to question what they were asked to build and consider the consequences of such achievements. Nowadays engineers are more confident in their social role and have learned to say “no” when the products are questionable or environmental damage may occur—the generation of space debris is one pertinent example. Hopefully, some “lessons learnt” guid- ance may ensue from the case studies and failure analyses recorded in this book. In 1986 engineers said “go” to the Challenger launch—other engineers said “no” but were over-ruled and the space shuttle exploded shortly after lift-off. It is only in hindsight that we understand that decision making can be extremely difficult, but such decisions must consider input from all engineering disciplines and the recognition of material properties is vital. A casual review of the Contents and Index will suggest to the reader that the subject matter is likely to be of interest not only to spacecraft engineers, but in the broader sense, to workers in quite different areas where metals, organic materials, composites, ceramics and glass are used under terrestrial conditions or within high vacuum systems. Advancements in technology always produce questions related to the reliability of new systems. Materials testing to agreed codes of practice have been shown to help maximise the reliability of new materials, pro- cesses, and applications. Metallography (or “materialography”) has led to an increased understanding of failure modes. Much emphasis of this book has been placed on failure analysis investigations. Each case must be developed in a logical manner—large-scale ix
- 13. (macroscopic) features are initially investigated, then the microscopic features of the materials involved. Test specimen or samples of spacecraft hardware must be meticulously prepared, then examined using both light and electron microscopy. It is amazing how these techniques have evolved and how the recording of images has progressed. The author and his metallurgist contemporaries may well remember early student days when contributions to reports were exquisitely detailed hand drawn micrographs or images captured on photographic plates. The digital revolution has now enabled all levels of detail to be recorded using super-resolution microscopes and the future seems to be heading towards 3-dimensional microscopy. In this book I have endeavoured to achieve a reasonable balance between general back- ground knowledge and in-depth technical information. An elementary understanding of metals and materials on the part of the reader is assumed. I have deliberately excluded a compre- hensive account of the techniques employed in modern materials laboratories (unless specifically related to unusual space material test methods). Many texts are available and cited in the Reference section. The Appendices have been extended and include many Tables related to: spacecraft materials’ properties; alloy comparisons as they may be procured in different countries; a simplified M&P management guideline for universities; and, examples of Declared Materials and Processes Lists. The space industry is a key sector in driving economic growth and creating new jobs. By 2030, the global space economy is predicted to be worth £400 billion per annum. At the time of writing, the European space manufacturing industry alone has an unprecedented overall turnover at £6 billion and a total direct employment of 38,000 persons. New spaceports will be established and spaceplanes are most likely to be the next generations’ means for transporting commercial and scientific payloads into orbit. Many future spacecraft engineers, space sci- entist and technologists, all specialists in their own fields, may be aghast that some funda- mental, ‘old-hat’ information is contained in this book. But it is the lessons-learnt scenarios that have brought us to where we are today. The industry is expanding and new employees need to learn from our past mistakes and, at least, understand why certain design rules exist. The wide acceptance of the previous book has been most welcome, and I hope the new changes and additions will also find approval by my colleagues in the space industry and others in the wider engineering community. Bosham, West Sussex Barrie D. Dunn December 2015 x Preface
- 14. Acknowledgments This book has been brought about by the blending of various published research and inves- tigation projects that I have undertaken as a metallurgist for the European Space Agency, from some written works of others and from personal friends. I am especially grateful to the late Dr. Jacques Dauphin my former Division Head at ESA who gave the encouragement to undertake the writing of the earlier book. He was a native of the French province of Lorraine, where the motto is ‘Qui s’y frotte s’y pique’ which loosely translates to ‘gather thistles, expect prick- les’—quite an apt maxim for those of us who have been involved with failure investigations. I also acknowledge the help received from my former ESA colleagues: Dr. Ton de Rooij, Jack Bosma, Guy Ramusat, Adrian Graham, David Collins and David Adams. Special thanks are also given to Dr. Ernst Semerad, Dr. A. Merstallinger, Grazyna Mozdzen and Markus Fink of the Aerospace and Advanced Composites GmbH (formally ARC), Wr. Neustadt, Austria, with whom I have had many years of professional collaboration. As previously stated, there has been a marked progress in this field of materials technology, resulting in significantly more citations to references in this Edition, but even so, the bibliographic information certainly is not complete. Where I have forgotten to cite a reference or credit an image I hope the author will forgive my oversight. I am also grateful to ESA and NASA for some of the illustrations used in the book. It should be noted that the opinions expressed in this book are those of the author and do not necessarily reflect the policy of the European Space Agency. Let me add a special note of thanks to my late wife, Hanneke, my son, Martin, and my daughter Harriet, for their patience through the spare-time hours that went into the making of the previous Edition. Also, to Anne for her unswerving support and help editing this present book. Stephen Hulcroft’s assistance at BlueFish Computer Services, Chichester is appreciated. I also wish to thank Clive Horwood, and the staff at Springer Praxis Books in Germany (Ms. Janet Sterritt) and India (Mr. Antony Raj Joseph and Ms. Sivajothi Ganesarathinam), for their assistance during the publication of this book. The author would like to thank all his colleagues and friends at the following organisations who kindly supplied new information, reference material and photographs: Torbjörn Lindblom, Celsius Materialteknik, Karlskoga, Sweden. Dr. Michael Osterman, The Centre for Advanced Life Cycle (CALCE), University of Maryland, MD, USA. S. Clément, Centre National d’Etudes Spatiales, Toulouse, France. Dr. H. Boving, Centre Suisse d’Electronique et de Microtechnique SA, Neuchâtel, Switzerland. H. Papenberg, DASA-ERNO Raumfahrttechnik GmbH (now Airbus Industries), Bremen, Germany. D. Bagley, ERA Technology, Leatherhead, UK. Dr. A. Feest, The Harwell Laboratory, Metals Technology Centre, Harwell, UK. W. Feuring, Heraeus GmbH, Hanau, Germany. Massimo Bonacci, High Technology Center (HTC), Foligno, Italy. xi
- 15. Poul Juul, Hytek, Aalborg, Denmark. Messrs G. Kudielka and W. Maier, IFE, Oberpfaffenhofen, Germany. Luca Moliterni and Gianluca Parodi, Italian Institute of Welding (IIS), Genoa, Italy. Norio Nemoto, Japan Aerospace Exploration Agency (JAXA), Tsukuba, Japan. Dr. Suman Shrestha, Keronite International Ltd., Haverhill, UK. P. Fletcher, Airbus (formally MMS-UK), Portsmouth, UK. Dr. Christopher Hunt, Martin Wickham and Ling Zou, The National Physics Laboratory, Teddington, UK. Dr. David Bernard, Nordson DAGE, Aylesbury, UK. Jo Wilson and Bob Hussey, RJ Technical Consultants, Juicq, France. Messrs Jörgen Svensson, U. Berg and Hans Ollfors, RUAG (formally Saab Ericsson Space), Gothenburg, Sweden. M.P. Hayes, The Spring Research and Manufacturers’ Association, Sheffield, UK. Ian Turner, Cathy Barnes and Malcolm Snowdon, Spur Electon Ltd., Havant, UK. Dr. R. Eckert, Standard Elektrik Lorenz, Stuttgart, Germany. Dr. P. von Rosenstiel, Stichting Geavanceerde Metaalkunde, Hengelo, The Netherlands. Luca Soli and Ulisse Di Marcantonio, Thales Alenia Space Italia, Milan, Italy. Dr. J.M. Motz, Thyssen Guss AG, Mülheim a.d. Ruhr, Germany. Stephen Kyle-Henney, TISICS Ltd., Farnborough, UK Bill Strachan and Dr. Asa Barber, The University of Portsmouth, Portsmouth, UK. K. Ring, Zentrum für Verbindungs Technik, Gilching, Germany. Robert Wm. Cooke, NASA—Johnson Space Center, Houston, TX, USA Pablo D. Torres, NASA—Marshall Space Flight Center, Huntsville, AL, USA Dr. Fabiola Brusciotti, Tecnalia, San Sebastian, Spain xii Acknowledgments
- 16. Contents 1 Introduction. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 2 Requirements for Spacecraft Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 2.1 General Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 2.2 Considerations for Materials and Processes . . . . . . . . . . . . . . . . . . . . . . 10 2.2.1 General Considerations During the Selection of Materials and Processes. . . . . . . . . . . . . . . . . . . . . . . . . . . 10 2.2.2 Some Futuristic Ideas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11 2.2.3 Some Basic Considerations Regarding Corrosion Prevention . . . 17 2.2.4 Space Project’s Phases and Management Events. . . . . . . . . . . . 20 2.3 The Effect of a Space Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . 22 2.4 Materials for Space Launch Vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . 28 2.5 Non-metallic Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38 2.5.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38 2.5.2 Classes of Non-metallic Materials. . . . . . . . . . . . . . . . . . . . . . 42 2.5.3 Novel Non-metallics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43 2.6 The Potential for Welding and Joining in a Space Environment . . . . . . . . 49 2.6.1 Background Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . 49 2.6.2 Potential Joining and Cutting Processes . . . . . . . . . . . . . . . . . . 50 2.6.3 Expectations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 3 The Integration of ‘Materials’ into Product Assurance Schemes . . . . . . . . . . 55 3.1 General Product Assurance and the Role of Materials . . . . . . . . . . . . . . . 55 3.1.1 Product Assurance Management . . . . . . . . . . . . . . . . . . . . . . . 55 3.1.2 Quality Assurance. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55 3.1.3 Reliability and Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57 3.1.4 Materials and Processes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59 3.1.5 Component Part Selection, and Procurement . . . . . . . . . . . . . . 61 3.1.6 Control of Ground-Handling Facilities. . . . . . . . . . . . . . . . . . . 63 3.2 The Materials Laboratory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66 3.2.1 Major Objectives of Laboratory . . . . . . . . . . . . . . . . . . . . . . . 66 3.2.2 Facilities and Instrumentation. . . . . . . . . . . . . . . . . . . . . . . . . 67 3.2.3 The Use of New Laboratory Techniques for NDT . . . . . . . . . . 85 3.2.4 Organic Chemistry and Environmental Test Laboratories . . . . . . 98 3.3 Preparation of Materials and Metallographic Evidence. . . . . . . . . . . . . . . 100 3.3.1 The Metallographer. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 3.3.2 Laboratory Records and Reports. . . . . . . . . . . . . . . . . . . . . . . 101 3.3.3 Report of Materials Data to Spacecraft Projects . . . . . . . . . . . . 101 3.3.4 Training of Materials Engineers and Laboratory Staff . . . . . . . . 103 3.3.5 Ethical Issues. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 xiii
- 17. 3.4 The Future for Materials Failure Investigations. . . . . . . . . . . . . . . . . . . . 104 3.4.1 The Larger Company . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 3.4.2 The Smaller Company. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 3.4.3 Product Liability. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 3.5 ‘Greener’ Spacecraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 3.6 The Potential for Recycling Electronic Waste. . . . . . . . . . . . . . . . . . . . . 111 3.6.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111 3.6.2 Elemental Distribution for Spacecraft Electronic Box . . . . . . . . 111 4 Spacecraft Manufacturing—Failure Prevention and the Application of Material Analysis and Metallography . . . . . . . . . . . . . . . . . . . . . . . . . . . 115 4.1 Sources of Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115 4.2 Drawings and Workmanship . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115 4.2.1 Design and Manufacturing Drawings. . . . . . . . . . . . . . . . . . . . 115 4.2.2 Workmanship Standards . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116 4.3 Mechanical Damage Revealed by Microstructure . . . . . . . . . . . . . . . . . . 122 4.4 Hydrogen Embrittlement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122 4.4.1 Interaction of Metal with Hydrogen . . . . . . . . . . . . . . . . . . . . 122 4.4.2 Hydrogen Embrittlement of Spring Steel . . . . . . . . . . . . . . . . . 123 4.4.3 Blistering of Plated Aluminium Alloy . . . . . . . . . . . . . . . . . . . 124 4.4.4 Examination for Titanium Hydride Precipitates. . . . . . . . . . . . . 125 4.4.5 Embrittlement of Copper . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127 4.4.6 Future Developments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128 4.5 General Corrosion Problems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128 4.5.1 Bimetallic Corrosion-Related Failures . . . . . . . . . . . . . . . . . . . 128 4.5.2 Corrosion Resistance of Anodic and Chemical Conversion Coatings on Al 2219 Alloy . . . . . . . . . . . . . . . . . . . . . . . . . . 132 4.5.3 Evaluation of Alodine Finishes on Common Spacecraft Aluminium Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134 4.5.4 Cleaning, Passivation, and Plating of Spacecraft Steels . . . . . . . 137 4.5.5 Launch Site Exposure and Corrosion. . . . . . . . . . . . . . . . . . . . 138 4.6 Stress-Corrosion Resistance of Metals. . . . . . . . . . . . . . . . . . . . . . . . . . 139 4.6.1 Stress-Corrosion Cracking . . . . . . . . . . . . . . . . . . . . . . . . . . . 139 4.6.2 SCC Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140 4.6.3 The Properties of Spring Materials . . . . . . . . . . . . . . . . . . . . . 144 4.6.4 Bearing Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148 4.7 Control of Printed Circuit Boards . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148 4.7.1 Chemical Composition of Tin-Lead from Microstructure . . . . . . 148 4.7.2 Grainy Solder Coverage on PCBs and the Effects of Rework. . . 150 4.7.3 Evaluation of Multilayer Board Internal Connections. . . . . . . . . 155 4.7.4 Flexible Circuits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 159 4.7.5 Hot-Air-Levelled Circuit Boards. . . . . . . . . . . . . . . . . . . . . . . 160 4.7.6 Solder Assembly of Component Packages onto Multilayer Boards with High Heat Capacity . . . . . . . . . . . . . . . . . . . . . . 161 4.8 Control of Composite Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161 4.8.1 Metal–Matrix Composites for Space Structures. . . . . . . . . . . . . 161 4.8.2 Composite Contact Devices . . . . . . . . . . . . . . . . . . . . . . . . . . 164 4.8.3 Fibre-Reinforced Plastic Composites . . . . . . . . . . . . . . . . . . . . 166 4.8.4 Fibre-Reinforced Glass Ceramics . . . . . . . . . . . . . . . . . . . . . . 170 4.8.5 Carbon–Carbon Composites. . . . . . . . . . . . . . . . . . . . . . . . . . 170 4.8.6 Metal Matrix Composites for Spacecraft Pressure Vessels . . . . . 172 xiv Contents
- 18. 4.9 Control of Capillary Screens . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172 4.10 Examination of Electroless Nickel Deposits . . . . . . . . . . . . . . . . . . . . . . 173 4.10.1 Microcracked Electroless Nickel . . . . . . . . . . . . . . . . . . . . . . . 173 4.10.2 Electroless Nickel Plating of Aluminium Electronic Housings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175 4.11 Control of Electroforming Processes . . . . . . . . . . . . . . . . . . . . . . . . . . . 176 4.12 Dip Brazing of Aluminium Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179 4.13 Considerations for the Assembly of Subsystems by Welding . . . . . . . . . . 181 4.13.1 General Welding Methods and Controls . . . . . . . . . . . . . . . . . 181 4.13.2 Electron Beam Welding. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184 4.13.3 Laser Beam Welding. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 185 4.13.4 Explosive Welding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 186 4.13.5 Welding of Aluminium–Lithium Alloys. . . . . . . . . . . . . . . . . . 187 4.13.6 Welding of Thermoplastics for Space Applications . . . . . . . . . . 188 4.14 Control of Power System Weldments . . . . . . . . . . . . . . . . . . . . . . . . . . 189 4.14.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189 4.14.2 Welded Solar Arrays. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189 4.14.3 Suitability of Welded Battery Cells . . . . . . . . . . . . . . . . . . . . . 193 4.15 Problems Associated with Residual Stresses in Weldments . . . . . . . . . . . 195 4.16 Electromagnetic Emission from TIG Welding Equipment . . . . . . . . . . . . 195 4.17 Titanium Aluminides for High-Temperature Applications . . . . . . . . . . . . 196 4.18 Shape-Memory Alloys for Spacecraft Devices . . . . . . . . . . . . . . . . . . . . 197 4.19 Foamed Aluminium for Damping Purposes . . . . . . . . . . . . . . . . . . . . . . 202 4.20 Superplastic Forming and Diffusion Bonding of Metals. . . . . . . . . . . . . . 203 4.20.1 Forming of Propellant Tanks . . . . . . . . . . . . . . . . . . . . . . . . . 203 4.20.2 Diffusion Bonding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 4.20.3 Superplastic Forming and Diffusion Bonding in One Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 4.21 Cleaning of Mechanical Parts. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 4.21.1 General Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 4.21.2 Metallic Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 4.21.3 Cleaning of Individual Parts. . . . . . . . . . . . . . . . . . . . . . . . . . 210 4.21.4 Cleaning of Metallurgically Joined Assemblies. . . . . . . . . . . . . 213 4.21.5 Maintenance of Cleanliness . . . . . . . . . . . . . . . . . . . . . . . . . . 216 4.21.6 Cleaning of Silicone Contamination . . . . . . . . . . . . . . . . . . . . 219 4.22 Novel Thermal Management Materials . . . . . . . . . . . . . . . . . . . . . . . . . 220 4.23 Cold Sprayed Coatings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 223 4.24 Advanced Plasma Electrolytic Oxidation Treatment for Aluminium, Magnesium and Titanium Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224 4.24.1 General Process . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224 4.24.2 Characteristics of PEO Coatings . . . . . . . . . . . . . . . . . . . . . . . 225 4.24.3 Applications. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 4.25 Joining by “Friction Stir” . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 231 4.25.1 Friction Stir Welding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 231 4.25.2 Friction Stud Welding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 234 4.26 Selective Brush Electroplating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 234 4.27 Control of Coatings and Bonded Items by Tape Testing . . . . . . . . . . . . . 237 4.28 The Application of EB Welding Machine for Reflow Brazing . . . . . . . . . 239 Contents xv
- 19. 5 Metallography Applied to Spacecraft Test Failures . . . . . . . . . . . . . . . . . . . 247 5.1 Application of Electron Microscope . . . . . . . . . . . . . . . . . . . . . . . . . . . 247 5.1.1 SEM Examination of Fracture Surfaces . . . . . . . . . . . . . . . . . . 247 5.1.2 TEM Examination of Metallic Failures . . . . . . . . . . . . . . . . . . 250 5.2 Fasteners. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 251 5.2.1 Spacecraft Fasteners . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 251 5.2.2 Fastener Failure Due to Forging Defect . . . . . . . . . . . . . . . . . . 254 5.2.3 Laps and Surface Irregularities in Threads . . . . . . . . . . . . . . . . 255 5.2.4 Hydrogen Embrittlement of Steel Fasteners . . . . . . . . . . . . . . . 255 5.2.5 Embrittlement of Titanium Alloys. . . . . . . . . . . . . . . . . . . . . . 255 5.2.6 Galvanic Corrosion of Fasteners . . . . . . . . . . . . . . . . . . . . . . . 257 5.2.7 Contamination and Organic Fastener Lubrication Systems . . . . . 257 5.2.8 Metallic Particle Generation . . . . . . . . . . . . . . . . . . . . . . . . . . 258 5.2.9 Quality Assurance Controls for Fasteners. . . . . . . . . . . . . . . . . 261 5.3 Thermal History from Microstructure . . . . . . . . . . . . . . . . . . . . . . . . . . 262 5.4 Effect of Inclusions Within the Microstructure of Explosively Deformed Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264 5.5 Degradation of Passive Thermal Control Systems . . . . . . . . . . . . . . . . . . 266 5.5.1 General Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 266 5.5.2 Low-Emissivity Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . 268 5.5.3 High-Absorption Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . 269 5.5.4 Rigid Optical Solar Reflectors . . . . . . . . . . . . . . . . . . . . . . . . 270 5.5.5 Flexible Second Surface Mirrors. . . . . . . . . . . . . . . . . . . . . . . 271 5.6 Sublimation of Metals. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 272 5.6.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 272 5.6.2 Sublimation of and Condensation of Cadmium and Zinc . . . . . . 274 5.6.3 Heater Sublimation Problem Associated with Thruster Motor . . . 276 5.6.4 Sublimation of Klystron Cathode-Heaters . . . . . . . . . . . . . . . . 276 5.6.5 Sublimation of Rhenium . . . . . . . . . . . . . . . . . . . . . . . . . . . . 278 5.7 Beryllium for Spacecraft Applications . . . . . . . . . . . . . . . . . . . . . . . . . . 280 5.7.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 280 5.7.2 Health and Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 281 5.7.3 Integrity of Machined Beryllium. . . . . . . . . . . . . . . . . . . . . . . 283 5.7.4 Thermal Cycling on Work-Hardened Beryllium . . . . . . . . . . . . 284 5.7.5 General Etching Solutions for Beryllium . . . . . . . . . . . . . . . . . 285 5.7.6 Investigation of Microcracked Thin-Foil Detector Windows . . . . 286 5.7.7 Aluminium-Beryllium Alloys . . . . . . . . . . . . . . . . . . . . . . . . . 288 5.8 Deactivation of Catalyst Particles for Hydrazine Decomposition . . . . . . . . 288 5.8.1 Testing Procedure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 288 5.8.2 Material Investigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 288 5.8.3 Mechanism of Particle Deactivation . . . . . . . . . . . . . . . . . . . . 290 5.9 Cathode Emitter Degradation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 291 5.10 Investigation of a Failed Spacecraft Antenna . . . . . . . . . . . . . . . . . . . . . 293 5.11 The Wear of Ball Bearings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 296 5.12 Cold Welding of Mechanisms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 304 5.12.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 304 5.12.2 Cold Welding Due to Cyclic, Impact Loading . . . . . . . . . . . . . 306 5.12.3 Cold-Welding Due to Fretting . . . . . . . . . . . . . . . . . . . . . . . . 307 5.13 Defective Black-Anodized Electrical Connector . . . . . . . . . . . . . . . . . . . 308 5.14 Contaminant Particles—Identification of Their Sources . . . . . . . . . . . . . . 309 xvi Contents
- 20. 5.15 Silicone Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 310 5.15.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 310 5.15.2 Contamination of Black-Anodized Finish. . . . . . . . . . . . . . . . . 311 5.15.3 Contamination of Invar Moulding Tool . . . . . . . . . . . . . . . . . . 312 5.15.4 Removal of Silicone Polymers . . . . . . . . . . . . . . . . . . . . . . . . 314 5.15.5 Contamination of Aluminium Tubes for Vacuum Pinch-Offs . . . 317 5.16 Magnetic Problems. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 317 5.17 Thermal Stress-Induced Dimensional Changes . . . . . . . . . . . . . . . . . . . . 319 5.17.1 General Problems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 319 5.17.2 Stress-Relaxation by Thermal Gradients. . . . . . . . . . . . . . . . . . 319 5.17.3 Thermally Induced Vibrations . . . . . . . . . . . . . . . . . . . . . . . . 321 5.18 Defects in Titanium Piece-Parts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 323 5.18.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 323 5.18.2 Alpha-Case Embrittlement . . . . . . . . . . . . . . . . . . . . . . . . . . . 323 5.18.3 Titanium Hydride Embrittlement. . . . . . . . . . . . . . . . . . . . . . . 324 5.19 Leaking Water Tank on Launcher. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 325 5.20 Compatibility of Liquid and Solid Propellants with Components and Subsystems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 326 6 Failure Analysis of Electrical Interconnections and Recommended Processes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 329 6.1 Material Problems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 329 6.2 Welded Lead Wire Interconnections . . . . . . . . . . . . . . . . . . . . . . . . . . . 329 6.3 ‘Purple Plague’ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 332 6.4 Mechanical Electrical Connections . . . . . . . . . . . . . . . . . . . . . . . . . . . . 337 6.4.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 337 6.4.2 Wire-Wrapped Connections . . . . . . . . . . . . . . . . . . . . . . . . . . 337 6.4.3 Crimped Joints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 339 6.5 Soldered Interconnections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 340 6.5.1 Introduction to Soldering . . . . . . . . . . . . . . . . . . . . . . . . . . . . 340 6.5.2 Inspection of Soldered Joints . . . . . . . . . . . . . . . . . . . . . . . . . 341 6.5.3 The Effect of Thermal Fatigue on Solder-Assembled Leaded Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344 6.5.4 Effect of Thermal Fatigue on Leadless Components . . . . . . . . . 351 6.5.5 The Effect of Thermal Fatigue on Semi-rigid Cable Connections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353 6.6 Problems Associated with Coatings for Soldering Applications . . . . . . . . 357 6.6.1 The Need for Coatings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 357 6.6.2 Surfaces that Can Be ‘Soldered To’ . . . . . . . . . . . . . . . . . . . . 357 6.6.3 Surfaces that Can Be ‘Soldered Through’ . . . . . . . . . . . . . . . . 359 6.7 The Use of Indium Solder Alloys. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 363 6.8 Wires and Cables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 369 6.8.1 Selection of Plated Finish on Copper Conductors . . . . . . . . . . . 369 6.8.2 Effect of Ageing on the Solderability of Tin-Plated and Silver-Plated Wires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 371 6.8.3 ‘Red Plague’ Corrosion of Silver-Plated Copper, and Plagues on Other Plated Stranded Wires . . . . . . . . . . . . . . 375 6.8.4 Manganin Wire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 379 6.8.5 High-Voltage Wires, Cables, and Connections . . . . . . . . . . . . . 380 6.8.6 Cold Welding of Stranded Wires and Cables . . . . . . . . . . . . . . 380 Contents xvii
- 21. 6.9 Problems Associated with Soldering Fluxes . . . . . . . . . . . . . . . . . . . . . . 380 6.9.1 Purpose of a Flux . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 380 6.9.2 Heat-Shrinkable Sleeves Containing Solder Preforms . . . . . . . . 381 6.9.3 Stress Corrosion of Component Lead Material . . . . . . . . . . . . . 383 6.9.4 Flux-Corrosion of Silver-Plated Stranded Wires . . . . . . . . . . . . 383 6.9.5 Selection of a Soldering Flux or a Solderable Finish . . . . . . . . . 386 6.9.6 Control of Galvanic Corrosion . . . . . . . . . . . . . . . . . . . . . . . . 389 6.9.7 Cleaning of Flux-Contaminated Surfaces . . . . . . . . . . . . . . . . . 389 6.9.8 Flux Residues, Their Ingress into Top-Coat of PCB Surfaces, and Bake Out After Cleaning. . . . . . . . . . . . . . . . . . . . . . . . . 391 6.9.9 Conductive Anodic Filament (CAF) Formation and Particulate Contamination . . . . . . . . . . . . . . . . . . . . . . . . 394 6.9.10 Potential Health Hazards in the Electronic Assembly Area. . . . . 398 6.10 Problems Associated with Brazing . . . . . . . . . . . . . . . . . . . . . . . . . . . . 399 6.10.1 Design Considerations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 399 6.10.2 Brazeability of Materials and Braze Alloy Compositions . . . . . . 400 6.10.3 Brazing Fluxes and Their Removal . . . . . . . . . . . . . . . . . . . . . 403 6.10.4 Atmospheres for Brazing . . . . . . . . . . . . . . . . . . . . . . . . . . . . 404 6.10.5 Safety Precautions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 405 6.10.6 Produce Assurance Applied to Brazing Operations . . . . . . . . . . 405 6.10.7 Inspection Criteria for Brazed Aluminium Alloy Waveguide-to-Flange Joints . . . . . . . . . . . . . . . . . . . . . . . . . . 406 6.11 Diffusion Soldering/Brazing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 408 6.12 Effects of Rework and Repair on Soldered Interconnections . . . . . . . . . . 408 6.12.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 408 6.12.2 Cosmetics of Solder Fillets . . . . . . . . . . . . . . . . . . . . . . . . . . 410 6.12.3 Effect of Rework Electronic Components . . . . . . . . . . . . . . . . 410 6.12.4 Effect of Rework on Plated-Through Holes . . . . . . . . . . . . . . . 410 6.12.5 Effect of Rework on Composition of Joint. . . . . . . . . . . . . . . . 412 6.12.6 Recuperation of Unsolderable PCBs and Component Leads . . . . 413 6.13 Electrical Conductive Adhesives. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 413 6.14 Training and Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 415 6.14.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 415 6.14.2 Certification for Electronic Assembly Techniques . . . . . . . . . . . 417 6.14.3 Understanding Process-Induced Failures and the Importance of Workshops. . . . . . . . . . . . . . . . . . . . . . 418 6.15 Verification of Surface-Mount Technology and Prevalent Failure Mechanisms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 419 6.15.1 Verification Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 419 6.15.2 Failure Under Mechanical Overloading . . . . . . . . . . . . . . . . . . 422 6.15.3 Failures Due to Board Flatness Problems. . . . . . . . . . . . . . . . . 422 6.15.4 Failure Due to Co-planarity Problems . . . . . . . . . . . . . . . . . . . 423 6.15.5 Solder Joint Failure Due to Thermal Mismatch Between SMD and Substrate . . . . . . . . . . . . . . . . . . . . . . . . . 425 6.15.6 Conductor Track Failure Due to Thermal Mismatch . . . . . . . . . 428 6.15.7 Failure of RF Cables Connected by SMT . . . . . . . . . . . . . . . . 428 6.15.8 SMT Solder Joint Failure Due to Conformal Coatings. . . . . . . . 428 6.15.9 SMT Problems Related to Flux and White Residues . . . . . . . . . 432 6.15.10 Area Grid Array (AGA) Packaging. . . . . . . . . . . . . . . . . . . . . 434 xviii Contents
- 22. 6.15.11 High Voltage Interconnections and Influence of Geometry (Workmanship) on Corona Discharge . . . . . . . . . . 442 6.15.12 Tin Pest. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 448 6.15.13 Mechanical and Electrical Properties of Electronic Materials at Temperatures Down to 4.2 K . . . . . . . . . . . . . . . . 451 7 Whisker Growths. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 461 7.1 The Problem of Whisker Growth . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 461 7.2 Analysis of Failures Due to Whisker Growth . . . . . . . . . . . . . . . . . . . . . 462 7.2.1 Molybdenum Whiskers on Metallized Miniature Circuits . . . . . . 462 7.2.2 Tungsten Whisker Growth Within Travelling Wave Tubes. . . . . 466 7.2.3 Metal Oxide Whisker Precipitation in Glass Seals. . . . . . . . . . . 466 7.2.4 Integrated Circuit Failure Modes Due to Electromigration—Aluminium Whisker Growth and Solder Joint Voiding . . . . . . . . . . . . . . . . . . . . . . . . . . . . 468 7.3 Tin Whisker Growths . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 472 7.3.1 Tin Whisker Growth on a Plated Steel Housing . . . . . . . . . . . . 472 7.3.2 Tin Whisker Growth on PCB and Other Electronic Materials During Thermal Cycling . . . . . . . . . . . . . . . . . . . . . 474 7.3.3 Tin Whisker Growth on Crimp Termination Devices. . . . . . . . . 479 7.3.4 The Nucleation, Growth and Mechanism of Growth of Tin Whiskers—Results from a C-Ring Test Programme . . . . . 481 7.3.5 Some Properties of Tin Whiskers . . . . . . . . . . . . . . . . . . . . . . 485 7.4 Precautions to Avoid General Whisker Growths . . . . . . . . . . . . . . . . . . . 491 7.5 The Creation of Lead-Free Control Plans. . . . . . . . . . . . . . . . . . . . . . . . 494 7.5.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 494 7.5.2 Methods for Reprocessing Pure Tin Terminations . . . . . . . . . . . 495 7.5.3 Mitigation Approaches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 498 8 Assessment of Post-flight Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 501 8.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 501 8.1.1 Hardware Return from Space . . . . . . . . . . . . . . . . . . . . . . . . . 501 8.1.2 Raw Materials from the Moon . . . . . . . . . . . . . . . . . . . . . . . . 501 8.1.3 Recent Investigations Using Retrieved Materials. . . . . . . . . . . . 503 8.2 Space Environmental Effects from Vacuum and Radiation. . . . . . . . . . . . 503 8.2.1 Organic Materials and Lubricants . . . . . . . . . . . . . . . . . . . . . . 503 8.2.2 Radiation Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 507 8.2.3 Effects of Vacuum on Metals. . . . . . . . . . . . . . . . . . . . . . . . . 508 8.3 Temperature Cycling. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 509 8.4 Micrometeoroids and Debris . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 509 8.4.1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 509 8.4.2 Debris Emanating from Catalytic Bed Thruster Motors . . . . . . . 512 8.4.3 Returned Hardware . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 514 8.4.4 Protection Shields. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 515 8.5 Effect of Atomic Oxygen on Materials . . . . . . . . . . . . . . . . . . . . . . . . . 517 8.6 Decelerators and Heat Shield Materials . . . . . . . . . . . . . . . . . . . . . . . . . 524 8.6.1 General Examples. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524 8.6.2 Beryllium as a Heat Shield . . . . . . . . . . . . . . . . . . . . . . . . . . 528 8.6.3 Alternative Heat Shield Materials . . . . . . . . . . . . . . . . . . . . . . 531 8.6.4 High-Temperature Fasteners. . . . . . . . . . . . . . . . . . . . . . . . . . 533 Contents xix
- 23. 8.7 Manned Compartments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 535 8.7.1 General Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 535 8.7.2 Solder Assembly Defects. . . . . . . . . . . . . . . . . . . . . . . . . . . . 538 8.7.3 Inspection of Spacelab Post-flight Hardware. . . . . . . . . . . . . . . 542 Appendix 1: Coefficient of (Linear) Thermal Expansion for Selected Materials (COE or CTE) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 557 Appendix 2: Properties of Printed Circuit Laminates . . . . . . . . . . . . . . . . . . . . 559 Appendix 3: Reagents for Microetching Metals and Alloys . . . . . . . . . . . . . . . . 561 Appendix 4: Conversion Table for Mechanical Properties . . . . . . . . . . . . . . . . . 565 Appendix 5: Aluminium Alloy Temper Designations . . . . . . . . . . . . . . . . . . . . . 567 Appendix 6: Metal Alloy Comparison Tables . . . . . . . . . . . . . . . . . . . . . . . . . . 571 Appendix 7: Variation of Standard Free Energy of Formation of Oxides with Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 613 Appendix 8: Simplied Procedure for the Management of Materials, Processes and Mechanical Parts—Possible Guidelines for a Cubesat or Small University Spacecraft . . . . . . . . . . . . . . . . 615 Appendix 9: Materials and Processes Standards Related to Space (Released by ECSS, JAXA and NASA) as of 2015 . . . . . . . . . . . . . 619 Appendix 10: Examples of Declared Process Lists (DPL). . . . . . . . . . . . . . . . . . 621 Appendix 11: Examples of Declared Materials Lists (DMLs) . . . . . . . . . . . . . . . 625 Glossary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 629 References. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 639 Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 655 xx Contents
- 24. 1 Introduction It is always impressive to look at the hardware of space ventures, whether in the form of launch vehicles culminating with the successful launch and landings of the Space Shuttle, of land-sited test chambers, of satellites being tested, of large antenna dishes, or of complex electronic circuitry under high magnification. But this is not the real space capability. The real capability lies in the people, in their technical compe- tence, and in their manipulations of the metals and materials which have made space communication programmes and space science programmes possible. Every 24 h an average of 45,000 storms break around the world. Meteorology is probably one of the oldest sciences, but it is presently one of the least accurate even though the four elements which enable the weather to be forecast (clouds, atmospheric pressure, temperature, and wind) are perfectly measurable. The various space programmes undertaken by the US, the USSR, Europe, and more recently Japan, India, China, South Korea and Brazil, are providing considerable amounts of information about the environment. The weather develops in a restricted area consisting mainly of the troposphere, the lowest layer of the atmosphere, which never exceeds 16 km in depth—barely twice the height of Mount Everest. But no part of the atmosphere will act independently of the other, and there is a continuing need to answer such questions as, how stable is our climate, or, how much additional atmospheric or water pollution can be tol- erated without drastically altering it? Such knowledge comes in many ways, and no more dramatically than in recent years from sensors in satellites orbiting the Earth, as illustrated in Figs. 1.1, 1.2, and 1.3, from observations and photographs made by astronauts, and even by correlation with the atmospheres of Mars, Venus, and Saturn which have also been investigated recently by space probes and landers. Wonderful images made by the Hubble Space Telescope have focused on galaxies in an infinite universe, but this orbiting observatory has also assisted in the expansion of knowledge of our own planet. The greenhouse effect, in which pollutants put into the atmosphere by the burning of fossil fuels and industrial processes trap heat and lead to the harmful warming of the planet, can be evaluated by using the new Earth resources satellites. These remote sensing spacecraft can also study deforestation, damage to the rain forests of South America, Africa, and Asia, and the effect of sewage and industrial waste on our oceans and shorelines. Meteorology, Earth phenomena observations, space-based satellite navigation systems (GPS) and com- munications by satellites are four major areas of more immediate interest where space technology can be applied usefully for the service of mankind. GPS saves mankind billions of dollars per year in wasted fuel for cars, aircraft and wear/tear! The Galileo global navigation satellite system (GNSS) should see 30 satellites in operation by 2019, pro- viding users with horizontal and vertical position measure- ments within a 1-m precision. Telecoms satellites in a geostationary orbit, 22,236 miles above the Earth’s surface enable internet signals to travel from one user to another anywhere in the world within 700 ms. These satellites allow us to speak and to see from one country or continent to another; to receive the same television picture simultane- ously in Britain, Luxemburg, America, Russia, or New Zealand will educate and entertain the people of the world— we may hope that the so-called telecom satellites will help to abolish the frontiers of misunderstanding and ignorance. This could be the key to the continuation of our civilization. The aforementioned attributes to space also support the United Nations’ goals of enhancing the literacy of everyone on the plant, eradicating poverty and improving health. Of particular concern to the space industries is the unpredictability of the Sun’s weather. It has been predicted that after years of relative calm, the years 2015 onwards may see a period of more intense activity such as solar flares, coronal mass ejections and magnetic storms. This so-called ‘space weather’ has been added to the UK’s National Risk Register as neutron-storms risk degrading satellite data and hardware, as well as on-Earth facilities that use sensitive electronic components such as metal-oxide-semiconductor © Springer International Publishing Switzerland 2016 B.D. Dunn, Materials and Processes, Springer Praxis Books, DOI 10.1007/978-3-319-23362-8_1 1
- 25. field-effect transistors (MOSFETs). Five spacecraft, in the project called High Energy Solar Physics Data in Europe, will continue to observe solar flares, explosions on the Sun and the magnetic phenomena that flings radiation and par- ticles across the Solar System. Permanently manned space stations such as the Rus- sian MIR and more substantially, the International Space Station—expected to remain operational until 2024—have been realized. Chinese scientists are planning an orbital space station where the core module will be launched in 2018, followed by two laboratory modules in 2020 and 2022. Settlements on the Moon are periodically proposed by NASA and other space agencies, but since the discovery of water-ice close to one of the lunar poles, this concept may become a reality. The remarkable Mars exploration projects involving orbiting spacecraft and rovers are establishing the Fig. 1.1 Example of visible image: clouds appear white and space is seen as black (courtesy NASA) 2 1 Introduction
- 26. “habitability” of that planet. Now that it has been established that water exists below the Martian surface, the future presence of manmade colonies on Mars may not be just the fantasy of science-fiction writers. The advancements in space technology have been made possible by many specific breakthroughs in materials science, manufacturing processes and novel technological advance- ments. Novel organic and inorganic materials have been developed by new methods of synthesis and metallurgical processing and these are facilitating the development of highly sophisticated spacecraft subsystems and electronic devices. As these developments have continued it is perhaps surprising to note that, in the main, it is the ‘ordinary’ metals and plastic materials and their related manufacturing or assembly processes that find their way into the construction of space hardware. Consequently we have now determined that these ‘ordinary’ materials produce the greatest prepon- derance of failures during the various stages of spacecraft Fig. 1.2 Example of infrared image of the Earth from space; cold and high clouds appear white. Warmer (low) clouds grey (courtesy ESA) 1 Introduction 3
- 27. programmes. The understanding of material and component limitations is a real requirement during the development of spacecraft systems in order that materials engineers can predict realistic margins of safety. Failures may occur as a result of over-testing, overload- ing, or over-pressurization when it can be demonstrated that no fault is attached to the material itself. Other failures occur because of poor choice of material, shortcomings in design, mistreatment during construction, or when the part was not adequate to withstand a particular fatigue or corrosive environment. Spacecraft failures occur during their fabrica- tion, assembly, integration, and environmental testing, and the generally short period of approximately four years between design and launch necessitates that failure analyses are rapidly performed to identify failure modes and their causes. This can be achieved only when sufficient informa- tion is available about the history of the failed part, from its initial composition and heat treatment, through manufac- turing details to records of all post-manufacture storage and testing. Fortunately, the high degree of surveillance by the Product and Quality Assurance teams of space hardware contractors usually enables proper documentation of most details pertinent to a failure. This knowledge of the precise operating stresses and environmental conditions is of great help in the diagnosis of a failure mode. There is a grave danger that the line of development of space equipment and instrumentation may be lost if care is not taken to preserve documentation related to past failures. Much information is contained in in-house laboratory reports which are often filed and forgotten. This book may go a little way to collate a small percentage of the examinations per- formed by the author in order that similar design or pro- duction problems may be minimized in the future. The classical failure modes of fatigue, stress corrosion cracking, hydrogen embrittlement, and the degradation of polymers by ageing, outgassing radiation, etc. have frequently been associated with spacecraft hardware. It is expected that chemists, metallurgists and other material or manufacturing engineers, who have metallic failures to contend with, will be able to draw parallel examples to their work from the illustrated case histories included within this book. Many of these examples necessitate that some of the material requirements for space flight are understood, as these are frequently directly attributable to particular failure modes. An overview of the specific requirements of spacecraft materials is given in Chap. 2, and the role of material Fig. 1.3 Example of water-vapour channel image. Taken by Meteosat Second Generation spacecraft in 2006 (courtesy Eumetsat) 4 1 Introduction
- 28. evaluations vis-à-vis spacecraft product assurance schemes is described in Chap. 3. The case histories have been selected from a large num- ber of examinations involving standard material testing techniques, and they are divided into three characteristic groupings. Chapter 4 concerns problems encountered during certain spacecraft manufacturing phases; Chap. 5 relates to failures which possibly occurred during testing; and Chap. 6 concerns failures which may cause deterioration of electrical interconnections. Chapter 7 deals with the many aspects of whisker growth as they might affect practising materials engineers associated with structural materials, electronics, and space-related industries. The case histories presented are limited to elec- trical device failures that have resulted directly from the growth of whiskers. Corrective actions are proposed for each case, and it will be evident that the published literature does not hold a solution for every situation. Chapter 8 is a short overview of the effect that the space environment has on materials. Here, examples are given from the large number of analyses made on materials returned from Low Earth Orbit (LEO). The effects of out- gassing, temperature, micrometeoroids, and atomic oxygen are illustrated from data amassed by Space Shuttle flight experiments, the Long Duration Exposure Facility, Eureca, Medet and materials retrieved from the Hubble Space Telescope during its various repair missions. The major tool utilized for each investigation is that of microscopical metallography—a technique developed by H. C. Sorby in the 1860s for the examination of geological samples. However, the number of tools available to materials engineers for the examination of polymeric molecules, microstructure and the like, has increased enormously over the last decade. It is often necessary to utilize advanced instruments such as ESCA and Auger spectrometers, laser microprobe analysers, high-definition radiographic units, scanning-laser acoustic microscopies, infrared spectroscopy, etc. as diagnostic tools for failure analysis. These are fre- quently available on loan, or can be rented from local uni- versities or research establishments. It must be emphasized that the majority of these tools give very limited information to the inexperienced. Only with sufficient practical experi- ence or on-the-spot guidance will the investigator be able to piece together information gained from various stages of nondestructive testing, superficial binocular microscope observations, physical and analytical tests, and metallography. In the scope of this book it is not possible to include a description of the equipments selected for particular inves- tigations, nor is there place for detailed accounts of the methods which have been chosen. The traditional guidelines applicable to all engineering failures have been followed by the author who is indebted to his highly professional col- leagues as well as access to well-equipped materials labo- ratories and test facilities. As will be discussed, the majority of organic material and metallurgical investigations are made during hardware production or after equipment level testing. Some defective items also originate from units or structures which have been installed on-board engineering or qualifi- cation model spacecraft, so that results and recommenda- tions from ‘material and failure review boards’ can be fed back to project designers and engineers. This procedure will, we hope, eliminate future problems with flight model spacecraft. This book deals not only with failure analysis but also with the measures which may be taken for failure pre- vention by improving product reliability. Finally, the term ‘failure’ can be construed to have many meanings ranging in scope from the trivial to the calamitous. Throughout this text ‘failure’ is employed as a technical term meaning cessation of function or usefulness. 1 Introduction 5
- 29. 2 Requirements for Spacecraft Materials 2.1 General Background The Space Age began in 1957, with an 83 kg Russian Sputnik satellite bleeping greetings to a surprised world. Since that spectacular beginning, intensive effort has gone into the scientific exploration of space, exploration of the Moon and distant planets, manufacturing of materials in space laboratories, and exploiting orbiting satellites for communication, navigation and observation of the Earth. The early steps have passed into history, and most equip- ment and instrumentation has been and will continue to be replaced by lighter and more complex substitutes. The remarkable achievements of the Apollo Lunar Exploration Programme two decades ago still tend to overshadow the unmanned automated satellite flights, and it is not always realized that spacecraft orbiting above all continents of the world have already revolutionized global communications, maritime navigation, and worldwide weather forecasting. These satellites are now vital links in a global network. They would not have been economically or technically feasible before the advent of near-Earth space explorations. Satellite communications started on a commercial basis with the launch of Early Bird in 1965, less than eight years after the launch of the first Sputnik. This was the first satellite to remain stationary over the Earth, and it was able to provide a continuous connection between any two Earth stations. Until comparatively recently these so-called ‘ap- plications’ satellites were merely assemblies of separately designed components rather than thoughtfully integrated systems. Often component interfaces failed to match, reducing the overall system performance. These satellites, and to a more limited extent the ‘scien- tific’ satellites, are now incorporating standardized subsys- tems in an attempt to optimize performance factors including weight, reliability, and cost. It seems likely that the spacecraft designer has placed greatest emphasis on mass, as this is usually set by the capabilities of the assigned launch vehicle which will take the satellite from the Earth’s surface and inject it into the desired orbit. The lighter the satellite, the cheaper will be the launch costs. Another major performance factor, reliability, can also be purchased if money is preferentially funneled into reliability and test programmes rather than launch vehicles. The important point is that performance factors of weight, reliability, and cost are all interrelated. The designer of an applications satellite will be more willing to pay for the reliability level that would give him 10 years of operation than the designer of a scientific satellite designed to shut off transmission after only one year when the mission objectives are attained. One of the major aims of the European satellite manu- facturer has been to set up a European communications programme which will develop and launch long-life satel- lites. A supporting technology programme has been under- taken to develop and qualify most of the critical subsystems that will enter the design of future operational satellites. An experimental satellite (Orbital Test Satellite—OTS) was launched in 1978 to evaluate and test the performance of the various subsystems of future European communication satellite systems. OTS and its launcher are illustrated in Figs. 2.1a and 2.2. Its major subsystems under evaluation included: • communications—to relay information (data and com- mands) between Earth and satellite and, in concept, to and from other spacecraft. • power supply—to provide electrical power to all satellite subsystems. • on-board propulsion—to provide thrust for orbit changes, station-keeping, and deorbiting. • Environmental control—to maintain specified tempera- tures, radiation levels, electromagnetic environment, etc. • structure—to support and maintain satellite configuration on the ground, during launch and in orbit. © Springer International Publishing Switzerland 2016 B.D. Dunn, Materials and Processes, Springer Praxis Books, DOI 10.1007/978-3-319-23362-8_2 7
- 30. Fig. 2.1 a OTS ‘structural model’ during vibration testing in 1975. Thermal blankets are not yet fitted. This is the first ESA communication satellite and has a height of 2.5 m (ESA). b View of the Alphasat satellite, after tests in the Intespace’s anechoic test chamber, Toulouse, France, 15 March 2013. This communications satellite is 7.1 m high (ESA). c These 9 m-high spike-lined walls enclose the hushed interior of ESA’s Maxwell test chamber, which isolates satellites from all external influences to assess their electromagnetic compatibility (ESA) 8 2 Requirements for Spacecraft Materials
- 31. The general development plan for a new satellite type such as OTS involves the building of several test models such as a structural model, thermal model, and engineering model (refurbished from the thermal model), before con- structing the qualification model and finally a flight spacecraft. Alphasat, shown in Fig. 2.1b is a high-power telecom satellite built by Astrium, through a public–private partner- ship between ESA and UK operator Inmarsat. It is based on the mighty Alphabus, the new European telecom platform developed by Astrium and Thales Alenia Space under joint contract from ESA and the French space agency, CNES. Alphabus is Europe’s response to increased market pressure for larger telecom payloads for direct-to-home TV broad- casting, digital audio broadcasting, broadband access and mobile services. Alphabus incorporates innovative tech- nologies including: • electric propulsion—to optimise the satellite’s mass in favour of payload • modular payload—including an antenna module which can be adapted for different missions • star trackers—ensure highly accurate attitude and orbit control • lithium ion cell batteries—charged from high-perfor- mance solar cells Whereas OTS generated 1260 W from its pair of solar panels, feeding to two 24 Ah NiCd batteries and had a weight of only 1490 kg, Alphasat can accommodate missions with up to 18 kW of payload power and has a weight of 6000 kg. Organisations such as the EU, ESA and NASA use mea- sures to assess the maturity of evolving technologies which can be related to devices, materials, components, etc. Regarding materials, mechanical parts and manufacturing processes, a new breakthrough or invention will not be suit- able for immediate use and some basic research will have to be conducted. This may lead to the technology being assessed for feasibility, for development and later the technology may be demonstrated in a laboratory environment. Validation of the new material may be made according to certain test methods. Mechanical parts may be qualified and manufacturing pro- cesses may be verified by the testing of “technology Fig. 2.2 a Launch of the European orbital test satellite (OTS-2, in 1978) on a Thor Delta rocket at Cape Canaveral. The 2 TV channels and 5000 telephone circuits operated without defects between 52 ground-stations (between Norway and Egypt) (ESA). b Launch of Alphasat—on the 25th July 2013, an Ariane 5 lifted off Europe’s largest telecommunications satellite (ESA) b 2.1 General Background 9
- 32. samples”—these are the steps usually taken in order to get approvals for space use by authorities (ECSS-Q-STD-70 2014). As a guide, the following listing can be used to assess the level of readiness of any materials technology: Technology Readiness Level Description TRL 1 Basic principles observed and reported TRL 2 Technology concept and/or application formulated TRL 3 Analytical and experimental critical function and/or charac- teristic proof-of-concept TRL 4 Component and/or breadboard validation in laboratory environment TRL 5 Component and/or breadboard validation in relevant environment TRL 6 System/subsystem model or prototype demonstration in a relevant environment (ground or space) TRL 7 System prototype demonstration in a space environment TRL 8 Actual system completed and “Flight qualified” through test and demonstration (ground or space) TRL 9 Actual system “Flight proven” through successful mission operations The reader may consider the above Technology Readiness Levels (TRLs) during the selection of materials and processes for a new application intended for use on board a spacecraft or even during the construction of a ground station (launch site). Obviously for any technology: the lower the TRL the more time and effort will be required before the approving authority can give authorisation for its incorporation into a space system. The concept of TRL’s will not be addressed during the following chapters of this book as every approval of a space material, mechanism and process will depend on the very precise requirements of a given space project. For some projects there may be an accepted higher level of risk involved during the selection of technologies. Low budget space flight experiments, providing they do not constitute a risk to the overall project, might choose to fly breadboard models that can give sufficient data return to university projects. At the other end of the spectrum, manned flight safety management will differentiate between “systems safety” and “payload safety”. Space systems safety will be a trade-off between complex project elements using flight proven technologies—here astronaut safety must be of paramount importance. Payload safety will consider the materials, mechanical parts and manufacturing processes and whether the payload is essential for flight operations and crew safety. Payloads and experiments can fail and not cause a risk to the astronauts. However, the materials from which they are manufactured will be of particular concern as these may operate beyond their intended temperatures; it is essential that these, usually non-metallic materials, do not release toxic substances by off-gassing, nor any fire hazard because of the flammable nature of the piece-part. 2.2 Considerations for Materials and Processes 2.2.1 General Considerations During the Selection of Materials and Processes The change of emphasis in Europe from building scientific satellites during the 1970’s with designed mission lives of one or two years to the production of a new generation of application satellites, which must be assured for periods of greater than twenty years in a somewhat hostile space environment, has necessitated that a greater effort is placed on confirming the reliability of many materials and tech- nologies which have previously been accepted as virtually fault-free. Additionally, the new modular approach and the drive to standardize subsystems for easy and economical adaptation for different satellite missions has led to long ground storage periods. This can cause material degradation problems, particularly the decay of liquid and solid fuels and the general corrosion of sensitive surfaces and even stress corrosion of structural elements. A listing of materials approved and utilized for the fabrication of a satellite such as the aforementioned Orbital Test Satellite in 1975 included at least 500 different organic and inorganic materials. Each was preliminarily approved for use in a given application, bear- ing in mind the environmental conditions it has been designed to withstand. The Declared Materials Lists asso- ciated with multipurpose space platforms for large telecommunications payloads, such as the 6.6 t Alphasat launched in 2013, involve more than 1000 different mate- rials. Until the late 1980s metallic materials have been the basic building materials of all satellites and launch vehicles with only a limited number of inroads from carbon fibre 10 2 Requirements for Spacecraft Materials
- 33. reinforced plastics (CFRP). Because of their exact alignment requirements some solar panels, dish antennas, and antenna platforms are fabricated from CFRP which, because of its small coefficient of expansion, will retain dimensional accuracy under the changing temperature conditions of an orbit (−160 to +180 °C). Launch vehicles, satellites, space probes and manned modules are predominantly built by industrial concerns engaged in aircraft manufacture (e.g. “prime contractors” such as Boeing, Airbus, Lockheed Martin, Alenia, Aerospatiale and Astrium). Because of this, designers will prefer to choose structural and mechanical parts from traditional metal alloys and composites, and will limit manufacturing to joining and finishing technologies which already exist in their respective plants. When com- pared to a mass production industry there is often little incentive to promote the use of advanced materials and alloys which may improve reliability and be weight-saving but will suffer the drawback of requiring costly fundamental testing and qualification before being incorporated into space hardware. Discussions between customers, prime contractors and their sub-tier suppliers involve contract requirement negoti- ations related to Materials and Processes issues and a con- siderable number of reviews will take into account such topics as design, materials selection, and fabrication processes. These are held throughout the various stages of every space project, from inception on the drawing board until the envi- ronmental testing and qualification of manufactured hardware prior to launch. However, it is not until the actual hardware is seen that one is struck by the results of cooperation between the many engineering disciplines. It is probable that the introduction of computer-aided design now means that spacecraft subassemblies and piece-parts are being fabricated to the closest tolerances ever achieved. The optimization of structural weight and the smaller design margins mean that a thorough knowledge of the materials selected for the appli- cation must be well established. This is particularly true for new, advanced materials, as the small design margins means there is no longer a reserve of strength built into the structure, as was the case for earlier spacecraft, to cover ignorance of design loads or stress intensities. The safety margins required of materials are real, but the over-conservative designs orig- inating from so-called ‘gloom factors’ or scatter in materials properties should be a thing of the past. To illustrate the accuracy demanded of modern machin- ing capabilities one can consider the unfortunate situation of the flawed primary mirror of the Hubble Space Telescope (HST). The prime objective of the HST mission was to obtain images of astronomical objects in approximately ten times sharper detail than that obtained by ground-based telescopes. The HST 2.4 m mirror was designed to be a precisely calculated hyperboloid. Although the mirror is actually smooth to a precision of 1/64 the wavelength of light (or one-millionth of an inch), a calculation error caused the mirror which was originally launched to have been fabricated with a curvature that was too shallow with a total centre-to-edge error of about 2 µm (or 1/50 thickness of a human hair). The result was that light rays hitting the mirror edges eventually made focus to a point that was slightly away from where light rays from the centre of the mirror focused: a defect called spherical aberration. The HST, delayed three years by the Challenger disaster, was launched in April 1990. Despite the flawed mirror, which rendered many of Hubble’s initial observations fuzzy, the new spaceborne telescope quickly demonstrated the advantages of an orbiting platform free from the interference of the Earth’s atmosphere. After the dramatic December 1993 repair mission, using astronauts from Space Shuttle Endeavour to correct the mirror and solar panel (see Sect. 8. 2) problems, Hubble began to demonstrate its full potential to peer into the universe. 2.2.2 Some Futuristic Ideas Advanced materials are finding more and more applications in new designs, and this is particularly true of reinforced polymers based on carbon or Kevlar fibres, clean materials (with low outgassing), and several new types of lightweight metal alloys. The microminiature electronic circuits so important for the relay of enormous volumes of data within a fraction of a second are also incorporating new materials with unique physical characteristics. Microdevices continue to be designed and prototyped (David 1996)—today these are termed micro-electro-mechanical systems (MEMS). Although many MEMS devices have been manufactured, to date, the only devices that have flown are accelerometers and gyroscopes (de Rooij 2009). In the USA, the JPL Centre for Space Microelectronics Technology has already produced a micro seismometer having a diameter of 12 mm and a ‘camera on a chip’ about the size of a fingernail. These kinds of advancement will certainly lead to smaller, lighter, and less costly spacecraft for the future. Even the so-called ‘nano satellite’, weighing about 1–2 kg, is thought to be feasible due to breakthroughs in small-scale engineering of MEMS. The cost of launching a satellite into LEO by the Space Shuttle was about £14,000 per kilogram and now, £5000 to £12,000 per kilogram when an ELV is selected. Costs to place a spacecraft into a geosynchronous transfer orbit (GTO) are estimated to be between and £20,000 per kilo- gram. Either launch vehicles should become less expensive, or satellites need redesigning to become far smaller and lighter so that multiple payloads (or even nanosats) can be launched simultaneously. A proliferation in the number of miniaturised satellites (often referred to as CubeSats) have been built by companies such as Clyde Space and SSTL 2.2 Considerations for Materials and Processes 11
- 34. (now part of Airbus) but probably the majority of those presently in orbit originate from schools and universities. These have low construction costs combined with fewer materials and processes requirements. Many have applica- tions beyond those of academic research or technology demonstration, and are used for Earth observation and defense purposes. The future will see more advanced manufacturing pro- cesses involved with the construction of space hardware— even traditional methods such as casting and forging will probably be to closer specification and under more highly inert atmospheres. The autoclave curing of composites will be done under clean conditions without the use of low volatile organic materials and any mold release agents will not contain silicones as they are difficult to remove prior to painting. Friction stir welding and FricRiveting can be envisaged for joining metals to thermoplastics; and laser materials processing will involve localized, intense heating of solid targets and components by laser, to achieve ultrafast, novel and economic joining and surface engineering. It used to be impossible to select a material without a full knowledge of how it might best be processed into a final piece-part—but today it seems that 3D printing has opened up a world of endless possibilities for designing and creating everything from a complex space mechanism to printed chocolates and foods for astronauts! 3D printing, also known as additive manufacturing (AM), produces three- dimensional items by “printing” them, layer-by-layer from raw ingredients consisting of powdered plastics, aluminium alloys, titanium alloys, low expansion alloys and other spacecraft materials. The most usual processing method is to introduce metal powder into a laser beam—a precise depo- sition of either sintered or melted powder is directed onto a flat table. The laser beam is controlled by CAD programs to raster across the sintered metal powder layer so consolidat- ing the deposit before another layer is added. The item is then built up, layer-by-layer to create a net-shaped part. This revolutionary rapid protyping process can now be used to create finalized spacecraft parts with a 40 % weight saving. Some improvements in the chemical purity of the powders used should increase yields to 100 % and, already, a 3D printer is in operation on the International Space station. This “first” 3D printing in space was performed in December 2014 by Butch Wilmore as part of a Zero gravity demon- stration—engineers up-linked a custom-made digital design file of a ratchet wrench to the 3D printer and produced a tool measuring 11.4 cm in length. This process will enable fragile items to be “manufactured in space” without the need to incorporate “robustness” and extra weight for surviving the shocks, vibration and mechanical loads encountered during launch. The deposition of materials can be so-called, func- tionally graded, permitting one face of the deposit to have totally different properties when compared to the opposite face. Similarly, custom compositions can be 3D printed so that undesirable compositions (possibly brittle intermetallic compounds, magnetic phases or corrodible compounds) in any binary or ternary phase diagram can be avoided by only depositing the useful compositions. The waste-, weight- and money-saving attributes of AM have already attracted manufacturers in all fields of advanced technology to incorporate AM into the production lines for their cus- tomized parts. These include heat-resistant bosses for turbine cases, large bearing housings, rocket engine injectors, landing gear support struts, and numerous spare parts. Creative endeavors, like that noted in Fig. 2.3, may even enable additive layer manufacturing (ALM) to print a hab- itable structure on the Moon (Redahan 2014). Figure 2.3a illustrates how functional habitation modules could be brought from Earth; the surface of the thin-walled inflatable structures would then be coated by 3-D printing a powder made entirely of regolith, having a particle size of around 200 µm, onto the thin-wall until a sufficiently large wall thickness could be built up to protect human space-workers from radiation and micrometeoroids. Regolith is the name given to lunar dust and this local resource has already been encountered by humans and analyzed with respect to particle size and composition (see Fig. 2.3b, c). Quantitative optical and electron-probe studies by the UK Institute of Geological Sciences (Simpson 1970) have shown that lunar samples can contain ilmenite, pyroxene, chrome-titanium spinel, troilite, native iron, iron-nickel alloy, and even native copper (as shown in Fig. 8.1). This concept for constructing a human outpost on the Moon using lunar soil, and ways to monitor the buildings’ progress from Earth by means of an industrial CCD camera positioned on the printer, have been described by Ceccanti (2010) and Colla (2014). Another rapid prototyping process that demonstrates great promise has been described by Maxwell et al. (2013). Known as Hyperbaric Pressure Laser Chemical Vapour Deposition (HP-LCVD), this rapid prototyping process incorporates a mixture of reactive gases into which laser beams penetrate for the growth of materials from atomic level to large structures by means of thermally- or photolytically-induced decomposition of the gaseous pre- cursor. Exploitation of this HP-LCVD process, as prescribed by Dynetics and the NASA Marshall Space Centre (Maxwell et al. 2013), may enable 3D rapid prototyping in-space for the fabrication of components, replacement parts and even nuclear thermal propulsion systems—by the use of precursor gases and raw materials found, often in abundance, within our own Solar System. Numerous advanced materials and manufacturing tech- niques will be individually described in Chap. 4. 12 2 Requirements for Spacecraft Materials
- 35. Fig. 2.3 a An artist’s impression of an igloo, built on the Moon by means of a 3D printer attached to the arm of the robotic vehicle seen to the right. Printing powder material is Moon dust (regolith), processed into a cell-like structure of high strength—the idea is to initially inflate a thin folded dome brought from Earth and protect it with a cellular shell using the 3D printer—the pressurised enclosure so sheilds astronauts from solar radiation, micrometerites and severe thermal cycling. Concept and illustration courtesy of ESA and architects Foster + Partners. b Apollo 11 astronaut’s footprint in lunar soil, made up of small, dust-like particles of regolith (courtesy NASA photo AS11-40-5877). c Grain size distribution of lunar soil from three different sites; about 50 % is greater than 100 µm from Heiken et al. (1974) 2.2 Considerations for Materials and Processes 13
- 36. Table 2.1 Static corrosion potential of metals and alloys (de Rooij 1989a) Material EMF Potential The metals having the greater negative EMF will tend to corrode and form oxides EMF between a calomel electrode and a 3.5 % NaCl water solution (V) Platinum +0.17 Carbon +0.15 Gold +0.15 Rhenium +0.08 Rhodium +0.05 Tantalum +0.04 Silver −0.03 Ag10Cu braze alloy −0.06 A286 (15Cr, 25Ni, Mo, Ti, V) passive −0.07 AISI 316 (18Cr, 13Ni, 2Mo, rem Fe) passive −0.07 AISI 321 (18Cr, 10Ni, 0.4Ti) passive −0.08 AISI 347 (18Cr, 12Ni, +Nb, rem Fe) passive −0.08 AISI 301 (17Cr, 7Ni) passive −0.09 AISI 304 (19Cr, 10Ni, rem Fe) passive −0.10 Hastelloy C (17Mo, 15Cr, 5W,6Fe, rem Ni) passive −0.10 Nichrome (80Ni, 20Cr) passive −0.10 Monel 60 (65Ni, 0.2Fe, 3.5Mn, 2Ti, 27Cu) −0.10 Inconel 92 (71Ni, 16Cr, 7Fe, 3Ti, 2Mn) passive −0.11 17-7PH stainless st. (17Cr, 7Ni, 1.1Al) passive −0.11 AISI 309 (23Cr, 13Ni) passive −0.11 Titanium −0.12 Monel 400 (32Cu, 2.5Fe, 2Mn, rem Ni) −0.12 CDA 442 (71Cu, 1Sn, 38Zn) −0.12 CD A 715 (70Cu, 30Ni) −0.12 Molybdenum −0.12 MP35N (Ni, 35Co, 2.0Cr, 10Mo) passive −0.15 CDA 510 (96Cu, 4Sn, P) phosphor bronze −0.16 AISI 420 (0.35C, 13Cr, rem Fe) passive −0.17 AISI 434 (0.12C, 17Cr, 1Mo, rem Fe) passive −0.17 Bismuth −0.17 Waspaloy (59Ni, 19.5Cr, 13.5Co, 4Mo) passive −0.17 Nickel passive −0.18 Monel 67 (67.5Cu, 31Ni, 0.3Ti, 0.5Fe) −0.18 Copper phosphorus (4.5P, rem Cu) −0.18 Copper phosphorus (8.5P, rem Cu) −0.19 Copper phosphorus (10.5P, rem Cu) −0.20 Copper −0.20 CDA 110 (electrolytic tough pitch) −0.20 CDA 172 (2Be, rem Cu) −0.20 Gold-germanium solder (12Ge, rem Au) −0.20 Copper-Gold (25Au, rem Cu) −0.20 AISI 440B (17Cr, 0.5Mo, rem Fe) passive −0.23 (continued) 14 2 Requirements for Spacecraft Materials
- 37. Table 2.1 (continued) Material EMF Potential The metals having the greater negative EMF will tend to corrode and form oxides EMF between a calomel electrode and a 3.5 % NaCl water solution (V) Ti6A14 V (6A1.4 V, rem Ti) −0.24 Silicon −0.24 Tungsten carbide (94WC, 6Co) −0.25 CDA 240 (80Cu, 20Zn) −0.25 CDA 220 (90Cu, 10Zn) −0.25 CDA 752 (65Cu, 18Ni, 17Zn) −0.25 CDA 180 (60Cu, 40Zn) −0.26 CDA 464 (60Cu, 1Sn, 39Zn) −0.26 CDA 270 (63Cu, 37Zn) −0.26 CDA 298 (52Cu, 48Zn) −0.27 Nichrome 80/20 (80Ni, 20Cr) active −0.27 CDA 521 (7Sn, rem Cu) −0.27 CuA110Fe (10A1.3Fe, rem Cu) −0.27 Armco 21-6-0 (22Cr, 12Ni, rem Fe) −0.27 Inconel 92 (71Ni,16Cr, 7Fe, 3Ti, 2Mn) active −0.28 CuA112 (12Al, rem Cu) −0.29 Niobium (1Zr, rem Nb) −0.30 Tungsten −0.30 Nickel active −0.30 Kovar, Nilo ‘K’ (29Ni, 17Co, rem Fe) −0.30 Chromium active −0.31 Cobalt −0.32 Nitinol (45Ti, 55Ni) −0.33 Invar (36Ni, rem Fe) −0.38 Cerrotric (42Sn, rem Bi) −0.39 SnAg4C3.5 solder −0.42 Sn95Ag4.9In0.1 solder −0.43 SnAg4 solder −0.46 SnAg5 solder −0.46 Tin −0.46 Sn10Sb solder −0.48 Indalloy no. 10 (75Pb.251n) solder −0.48 Indalloy no. 7 (50Pb,501n) solder −0.49 Lead −0.50 Sn63 (63Sn, 37Pb) solder −0.51 Sn60 (60Sn, 40Pb) solder −0.51 Sn62Ag2 (62Sn, 36Pb, 2Ag) solder −0.51 Sn59Sb2 (59Sn, 39Pb, 2Sb) solder −0.51 Sn60Sb5 (60Sn, 35Pb, 5Sb) solder −0.51 Sn60Sb10 (60Sn, 30Pb, 10Sb) solder −0.51 Sn60Pb39.5Cu0.12P0.9 solder −0.51 PbSn5Agl.5 solder −0.51 Mild steel −0.52 (continued) 2.2 Considerations for Materials and Processes 15
- 38. Table 2.1 (continued) Material EMF Potential The metals having the greater negative EMF will tend to corrode and form oxides EMF between a calomel electrode and a 3.5 % NaCl water solution (V) AISI 304 (19Cr, 10Ni, rem Fe) active −0.52 AISI 420 (0.35C, 13Cr, rem Fe) active −0.52 AA 2219-T3.T4 (6.3Cu, 0.3Mn, 0.18Zr, 0.1V, 0.06Ti, rem Al) −0.56 AISI 440B (17Cr, 0.5Mo, rem Fe) active −0.59 AA 2014-T4 (4.5Cu, 1Fe, 1Si, 0.15Ti, rem Al) −0.61 AA 2017-T4 (4Cu,1Fe,1 Mg,0.1Cr,rem Al) −0.61 AA 2024-T3 (4.5Cu, 1.5 Mg, 0.6Mn, rem Al) −0.62 AA B295.0-T6 (2.5Si, 1.2Fe, 4.5Cu, rem Al) casting −0.63 In75Pb25 solder −0.64 Indalloy No. 1 (50In, 50Sn) solder −0.65 AA 380.0-F (8.5Si, 2Fe, 3.5Cu, rem Al) casting −0.66 AA 319.0-F (6Si, 1Fe, 3.5Cu, rem Al) casting −0.66 AA 333-0-F (9Si.1Fe, 3.5Cu, rem Al) casting −0.66 Indium −0.67 AA 2014-T6 (4.5Cu, 1Fe, 1Si, 0.15Ti, rem Al) −0.69 Cadmium −0.70 AA 2024-T81 (4.5Cu, 1.5Mg, 0.6Mn, rem Al) −0.71 AA 2219-T6,T8 (6.3Cu, 0.3Mn0.18Zr, 0.1V, 0.06Ti, rem Al) −0.72 AA 6061-T4 (1 Mg, 0.6Si, 0.25Cu, 0.2Cr, rem Al) −0.72 AA 4043 (12Si, 1Cu, 1Mg, rem Al) −0.74 AA 6151 (1 Mg, 1Fe, 0.25Sn, 0.15Ti, rem Al) −0.74 AA 7075-T6 (5.6Zn, 2.5Mg, 1.6Cu, 0.3Cru.03Cr, rem Al) −0.74 AA 7178-T6 (6.8Zn, 32Mg, 2Cu, 0.2Ti, rem Al) −0.74 AA 1160 (98.4Al) −0.75 Aluminium −0.75 AA 5356 (5Zn, 0.1Ti, 0.1Cr, rem Al) −0.75 AA 5554 (5 Mg, 1Mn, 0.25Zn, 0.2Cr, rem Al) −0.75 AA 1050 (99.5Al) −0.75 Al-3Li −0.75 AA 1100 (99.0Al) −0.75 AA 3003 (1.2Mn, rem Al) −0.75 AA 6151 (1Mg, 1Fe,0.8Mn, 0.25Zn, 0.15Ti, rem Al) −0.75 AA 6053 (1.3Mg, 0.5Si, 0.35Cr, rem Al) −0.75 AA 6061-T6 (1Mg, 0.6Si, 0.25Cu, 0.2Cr, rem Al) −0.75 AA 6063 (0.7Mg, 0.4Si, rem Al) −0.75 Alclad 2014 (4.5Cu, 1Fe, 1Si, 0.15Ti, rem Al) −0.75 Alcald 2024 (4.5Cu, 1.5Mg, 0.1Cr, rem Al, Al-clad) −0.75 AA 3004 (1.5Mn, rem Al) −0.76 AA 1060 (99.6Al) −0.76 AA 5050 (1.5Mg, rem Al) −0.76 AA 7075-T73 (5.6Zn,2.5 Mg,1.6Cu,0.3Cr,rem Al) –0.76 AA 5052 (2.5Mg, 0.25Cr, rem Al) –0.77 AA 5086 (4Mg, 0.5Mn, rem Al) –0.77 (continued) 16 2 Requirements for Spacecraft Materials
- 39. 2.2.3 Some Basic Considerations Regarding Corrosion Prevention It is necessary to ensure that any newly selected material will retain its functional properties during all stages of the spacecraft’s designed life, up to the end of the mission. During manufacturing, the material must not degrade because of contamination from processing steps such as the release agents used for items moulded from CFRP, or by cutting oils used in the machining of alloys. Galvanic and general surface corrosion must be avoided during environ- mental testing and ground storage by the correct selection of surface finishes such as anodic films, chemical conversion films, and paints. When electrical grounding is required, only contacts having a compatible coupling of less than 0.5 V should be chosen. The static corrosion potential for a large number of metals and alloys has been established (de Rooij 1989a) and is presented in Table 2.1. However, de Rooij has simplified this Table into Groupings of metallic alloys and the modified tabulation now appears as shown in Table 2.2. The material may need a high resistance to Stress Corrosion Cracking (SCC) before launch, and in such cases can be selected from those alloys listed in Table 2.3. The primary and secondary structures will be made from light alloys based on aluminium and magnesium, together with titanium and to a very limited extent beryllium. Nickel alloys are often selected for their high-temperature perfor- mance and oxidation resistance; they are often known by trade names, rather than by their specification code numbers. Commercially pure nickel, easy to form into complex shapes, is used in the construction of spacecraft electronics where its electrical and magnetic properties are crucial. Mechanical designers often select Inconel alloys 600 and 625 because they appear in Table 2.3, but it has recently Table 2.1 (continued) Material EMF Potential The metals having the greater negative EMF will tend to corrode and form oxides EMF between a calomel electrode and a 3.5 % NaCl water solution (V) AA 5154 (3.Mg, 0.25Cr, rem Al) –0.78 AA 5454 (2.8Mg, 1Mn, 0.2Ti, 0.1Cu, 0.2Cr, rem Al) –0.78 AA 4047 (12Si, rem Al) –0.78 Al-C −0.78 AA 5056 (5.2Mg, 0.1Mn, 0.1Cr, rem Al) –0.79 AA 7079–T6 (4.3Zn, 3.3Mg, 0.6Cu, 0.2Mn, 0.2Cr, rem Al) –0.79 AA 5456 (5Mg, 0.7Mn, 0.15Cu, 0.15Cr, rem Al) –0.79 AA 5083 (4.5Mg, 0.7Mn, rem Al) –0.79 AA 7072 (1Zn, 0.5Si, 0.3Cr, rem Al) –0.87 Beryllium −0.97 Zinc −1.03 Manganese −1.21 Erbium −1.34 Electron (4Zn, 0.7Zr, rem Mg) –1.55 ZW3 (3Zn, 0.5Zr, rem Mg) –1.57 AZ61 (6Al, 1Zn, 0.3Mn, rem Mg) –1.57 AZ31B (3Al, 1Zn, rem Mg) –1.60 Magnesium −1.60 HK31A (0.7Zr, 3Th, rem Mg) –1.61 Notes to Table 2.1 Compatible material couples are considered to have a maximum potential difference of 0.25 V for non-cleanroom environments 0.50 V for cleanroom or hermetically sealed environments This galvanic series chart is useful for a first approximation in selecting materials for corrosion control, but may be too simplistic for further dependence. It provides no information concerning corrosion rates or what will happen when three or more metals are electrically coupled. Service conditions such as ionic concentration, aeration, metal purity, etc. can change relative positions. Reversals, especially with metals that are very close in the series, such as steel and aluminium, can cause serious service problems, and specialized polarization studies are then recommended The majority of alloys present in this Table can be referred to in Appendix 6 which lists specification number, composition limits and equivalent British, French, German and US standards 2.2 Considerations for Materials and Processes 17
- 40. Table 2.2 Suggested compatible couples for bimetallic contacts (after de Rooij, based on Table 2.1) Key 0—Can be used without restriction 1—Can be used in a non-controlled environment (e.g. assembly area and general non-clean room environmnet) 2—Can be used in a clean room environment 3—Need specific measures to avoid corrosion when these combinations are selected 18 2 Requirements for Spacecraft Materials
- 41. Table 2.3 Alloys with high resistance to stress-corrosion cracking Steel alloys Alloy Condition Carbon steel (1000 series) Below 180 ksi UTS Low alloy steel (4130, 4340, D6AC, etc.) Below 180 ksi UTS Music wire (ASTM 228) Cold drawn 1095 spring steel Tempered HY 80 steel Tempered HY 130 steel Tempered HY 140 steel Tempered 200 series stainless steel (unsensitized) All 300 series stainless steel (unsensitized)a All 400 series ferritic stainless steel (404, 430, 444, etc.) All Nitronic 32 Annealed Nitronic 33b Annealed Nitronic 40 (formerly 21-6-9)b Annealed A-286 stainless steel All AM-350 stainless steel SCT 1000 and above AM-350 stainless steel SCT 1000 and above AM-362 (Almar 362) stainless steel 3 h. at 1000 °F Carpenter 20Cb-3 stainless steel All Carpenter 20Cb-3 stainless steel All Custom 450 stainless steel H1000 and above Custom 455 stainless steel H1000 and above 15-5PH stainless steel H1000 and above PH15-7Mo stainless steel CH900 17-7PH stainless steel CH900 Aluminium alloys Wrought Cast Alloyc Temperd Alloye Temper 1000 series All 319.0, A319.0 As cast 2011 T8 333.0, A333.0 As cast 2024 rod, bar T8 355.0, C355.0 T6 2219 T6, T8 356.0, A356.0 All 2418 T8 2618 T6 357.0 All 3000 series All B358.0 (Tens-50) All 5000 series Allf, g 359.0 All 6000 series All 380.0, A380.0 As cast 7049 T73 514.0 (214) As castg 7149 T73 518.0 (218) As castg 7050 T73 535.0 (Almag.35) As castg (continued) Table 2.3 (continued) Aluminium alloys 7050 T73 A712.0, C712.0 As cast 7475 T73 Copper alloys CDA No.h Condition (% cold rolled)i 110 37 170 AT, HT.j 172 AT, HT.j 194 37 195 90 230 40 280 0 422 37 443 10 510 37 521 37 524 0 606 0 619 40 (9 % B phase) 619 40 (95 % B phase) 655 0 688 40 704 0 706 50 710 0 715 0 725 50 Annealed Nickel alloys Alloy Conditions Glass Seal 52 CR (51Ni-49Fe) All Invar 36 (36Ni-64Fe) All Hastelloy B Solution heat treated Hastelloy C All Hastelloy X All Incoloy 800 All Incoloy 901 All Incoloy 903 All Inconel 600 Annealed Inconel 625 Annealed Inconel 718 All Inconel X-750 All Monel K-500 All Ni-Span-C 902 All Rene 41 All (continued) 2.2 Considerations for Materials and Processes 19
- 42. been acknowledged that these alloys soften, and can suffer from SCC in pure water, at temperatures above 300 °C (steam generators for nuclear plants). It may be wiser to use a development of the 625 alloy when high temperatures may be expected, for instance in propulsion systems. This development involved the additions of molybdenum and niobium to 625 to impart solid solution hardening and the formation of Ni3Nb, a very effective hardening precipitate. This is known as the super alloy Inconel 718 and has become the most widely used high temperature nickel alloy. All the classical assembly methods are employed: weld- ing, brazing, soldering, riveting, bolting, and adhesive bonding. It is important to ensure the joining processes themselves have not degraded the materials’ surface or stress corrosion resistance. (Heating can modify an alloy’s microstructure, weld metal and heat affected zones will be different to the parent metal, braze metal may be noble to the remaining surfaces which can preferentially corrode, mechanical joints can have re-entrant faces that retain water and cause pitting, even cured resins may release acids that damage the surrounding surfaces). Generally, aircraft industry manufacturing standards are followed, and much attention is given to process control and there is a need to evaluate all process used to join together structural and electrical parts. 2.2.4 Space Project’s Phases and Management Events It is important to note that before a satellite becomes fully operational in orbit its subsystems, mechanisms, and elec- tronics will have been subjected to the following main environmental conditions: (A) Ground activities: Operation for test and checkout Handling Transportation Storage Exposure to the elements (B) Subjection to launch and ascent: Acceleration and shock Vibration and acoustic noise and possibly contact with reactive fuels, oxidizers, and temperature extremes Pyrotechnic shock (C) Transfer to operational orbit position: Thermal cycling due to exposure to the Sun and eclipse Ultrahigh vacuum Radiation—electromagnetic and penetrating particles Zero gravity The new millennium saw a great increase in the size and complexity of spacecraft. This has necessitated ground test- ing facilities to become modernized, physically larger, and more sophisticated for the exposure of space hardware to the environments listed in A to C above. The Test Facility within space agencies are comprised of high capital investments such as Large Space Simulators that reproduce the vacuum, certain radiations and the cryogenic-to-elevated temperatures encountered by space hardware. Mechanical and acoustic tests simulate the launch environment, the magnetic Table 2.3 (continued) Nickel alloys Unitemp 212 All Waspaloy All Miscellaneous alloys Alloy Conditions Beryllium S-200C Annealed HS25 (L605) All HS 188 All MP35 N Cold worked and aged MP159 Cold worked and aged Titanium 3Al-2.5 V All Titanium 5Al-2.5SN All Titanium 6Al-4 V All Titanium 10Fe-2 V-3Al All Titanium 13V-11Cr-3Al All Titanium IMI 550 Al Magnesium MIA All Magnesium LA141 Stabilized Magnesium LAZ933 All a Including weldments of 304L, 316L, 321 and 347 b Including weldments c Including weldments of the weldable alloys d Mechanically stress relieved (TX5X or TX5XX) where specified e The former designation is shown in parentheses where significantly different. See Appendix 5 for temper designations f High magnesium alloys 5456, 5083, and 5086 should be used in controlled tempers (H111, H112, H116, H117, H323, H343) for resistance to SCC and exfoliation g Alloys with magnesium content greater than 3.0 percent are not recommended for high-temperature application, 66°C (150°F) and above h Copper Development Association alloy number i Maximum percent cold rolled for which SCC data are available j AT—annealed and precipitation hardened HT—work hardened and precipitation hardened Notes to Table 2.3 Data are compiled from NASA MSFC Spec 522B and ECSS-Q-ST-70-36. Recent issues of these documents should be consulted for classification of alloys with both a moderate and a low resistance to stress corrosion cracking. Appendix 5 describes the aluminium alloy temper designations. A search through Appendix 6 will indicate similar European alloys 20 2 Requirements for Spacecraft Materials
- 43. characteristics of spacecraft are evaluated and all tests are performed under various classes of cleanroom conditions. As mentioned previously, a development plan for a new spacecraft design will involve a ‘model philosophy’ where models of the craft will be dynamically tested—without testing the risk of failures is too great. In the early days it was not uncommon for space authorities to build four models for testing prior to actually building a flight spacecraft. The model philosophy will be accounted for in the next para- graph, but it is emphasised that as the space industry has matured, the design margins have become established (for structures and electronic systems) so that there is now more focus on analysis and less on actual testing. It is now the norm to build only one prototype for testing—often this ‘build and test’ is completed only six months before work commences on manufacturing the flight spacecraft. It is interesting to remember the aims of the ‘model philosophy’ for the early (70’s and 80’s) European telecommunication satellite projects. The structural model was subjected to a programme of tests which exceeded the expected launch environment conditions (each type of launch vehicle has its own characteristic levels of vibration and acoustic noise). Typical test configurations are shown in Fig. 2.1a–c. The enormous amount of energy released during launch can be witnessed from Fig. 2.2. Weakness in designs may be exposed by this model, such as failures resulting from fatigue of welds, struts, electronic box hold-down points, and the like. The thermal model was subjected to solar simulation and thermal balance testing to confirm and update previously determined mathematical models. During these tests, deficient designs may promote several material failures related to thermal fatigue, overheating, and embrit- tlement of incompatible joining techniques and metal alloys. By simply modifying the paint finish of the spacecraft sur- face, or by the attachment of reflective mirrors, it was found possible to adjust and reduce the local temperature envi- ronment of each subsystem or equipment and reduce the chance of thermally induced failures. Workmanship prob- lems abounded, with non-conformances relating to open circuits in cable harnesses due to wires separating from crimp barrels and cold soldered joints on circuit boards (such events became less frequent once operator and inspector training schemes were introduced—see Sect. 6.14). The engineering model ensured that integration and performance could be achieved. A review of the old project and labora- tory failure investigation reports was made by the author (at the time of filing and archiving as a result of which they were lost forever!). This revealed that several material problems only came to light during integration, particularly at the mechanical interfaces between equipments and the structure (e.g. failure of springs and bolting devices due to incorrect plating processes which cause delayed failure by hydrogen embrittlement; the over-torqueing and fracture of lock-nuts and other operator handling errors, etc.). It was seen that the reliability of electrical interfaces between equipments, and the mutual compatibility between the constituent subsys- tems, was seriously jeopardized by fundamental oversights (e.g. high electrical resistance between gold-plated and aluminium-finished interfaces due to galvanic corrosion; migration of silver to produce short circuits; the use of austenitic steels having work-hardened and therefore slightly magnetic surfaces in locations required to be magnetically clean; etc.). Remedial actions were taken by suppliers and assemblers as a result of failure review boards (FRBs). Lessons learnt documents (nowadays called internal problem notification documents—IPNs) were written and circulated. Finally, a fully assembled qualification model was built and subjected to a comprehensive series of environmental ground tests. These usually included: sine vibration, spin, acoustic noise vibration, centrifuge acceleration, and solar simulation. Each of these major test phases was preceded by an integrated systems test in order to verify that the func- tional behaviour of the satellite was correct. It was not uncommon that the qualification model would be recognized as a “flight spare” in case of launcher failure. Flight model test programmes were, and continue to be, more limited than those used on the qualification model (most customers now refer to this as the protoflight model). It may be assumed that some subsystem parts that will operate for relatively short periods after launch can accu- mulate several hundred hours of test operation before the actual time of launching. Materials selected for use under the vacuum conditions of space may therefore have to operate for periods under normal atmospheric conditions. This may create special problems. One is immediately reminded that very thin (tens of Ångstroms1 ) films of lead, or molybdenum disulphide, for the lubrication mechanisms, can rapidly oxi- dize under terrestrial conditions and become the cause of malfunction. This goes to illustrate the need to know the effects that ground testing may have on delicate surfaces. A final concern of the writer relates to the participation of materials experts on spacecraft project review boards. It is paramount that an experienced materials engineer is incor- porated into each of the four major reviews during the design and construction of individual spacecraft. Whenever ECSS Q-ST-70 is included as a contractual document this become a requirement—as a minimum, the materials engineer will manage the steps taken for the project-approval of declared lists for every flight material, mechanical part and their related processes (i.e. DML, DMPL and DPL). Tasks should also include cleanliness and contamination control, the testing and validation of new materials, assistance in the 1 1 Å equals 10−10 m. 2.2 Considerations for Materials and Processes 21
- 44. qualification of mechanical parts and the verification of new or critical processes. Spacecraft, such as CubeSats and university flight experiments may follow less rigorous requirements and a reduced M&P programme is suggested in Appendix 8. Reviews are usually contractual milestones and essen- tially question: will the design, hardware, software, and operational approach satisfy the mission objectives? The review names and their main objectives are as follows: PDR (preliminary design review): after evaluation of thermal and/or engineering models, to approve and release the preliminary design, including materials and processes. CDR (critical design review): this establishes the final design and agrees that flight hardware manufacturing can com- mence (all declared materials being approved and pro- duction process, when required, are verified as being suitable). QR (qualification review): assess that all qualification activities on subsystems are complete—for certain projects a qualification model is built. LRR (launch and operations readiness review): this checks that all lower-level acceptance reviews have been suc- cessfully completed, the flight model spacecraft accepted, and it is then authorized to be launched. 2.3 The Effect of a Space Environment (a) General The purpose of this section is to provide the reader with an overview of the salient points concerning the effect of a space environment on spacecraft parts and materials. Far greater discussion with examples will be made of these effects throughout the remainder of this book. Each spacecraft material will be required to suffer no vibrational fatigue damage during the launch. In orbit, it will need to survive the space environment (see Table 2.4) and, in particular, to possess low-outgassing- under-vacuum properties, whether it be a lubricating grease or a structural plastic. Radiation and thermal cycling must not degrade thermal-control surfaces or joints in materials possessing different coefficient of expansion. The presence of atomic oxygen in low Earth orbits, a relative newcomer in environmental effects, has been seen to lead to the erosion/corrosion/oxidation of many material surfaces, and more coatings with good atomic oxygen durability need to be developed. These aspects will be detailed in Chap. 8. (b) Sublimation and evaporation The minimum altitude for an Earth-orbiting satellite is 200 km (125 miles), and, once this altitude has been reached, appreciable changes can be produced in common engineer- ing materials, whether they be metals, plastics, or ceramics. The vacuum in space is very high, the pressure falling from 10−6 mm Hg at 200 km to less than 10−12 mm Hg beyond 6500 km. Some polymers will decompose and some metals will tend to sublimate under vacuum. The rate at which the molecules or atoms leave a surface in vacuum will rise rapidly with an increase in temperature according to the equation: G ¼ 5:04 103 P M=T ð Þ1=2 where G grams of material evaporated or sublimated per square centimetre per day P Vapour pressure of the evaporating species in torr2 M Molecular weight of the material T Absolute temperature, K Temperature has an enormous effect on the amount of metallic material that is sublimated. As examples, cadmium and zinc have sublimation rates that increase by a factor of ten for roughly every 30 °C rise in temperature: Cadmium has a vapour pressure of 10−8 at 70 °C (approximately) 10−7 at 90 °C 10−6 at 120 °C 10−5 at 150 °C and 10−4 at 180 °C The relationship between the vapour pressure of a metal (P) and temperature (T) is given by the following equation: P ¼ P/eE=RT where P∝ = a constant (i.e. vapour pressure at T = ∝), E = heat of evaporation (e.g. joule · mole−1 ), and R = gas constant (8.3 J mol−1 K−1 ). The temperatures for given metallic sublimation rates are listed in Table 2.5. The thermal environment in space is completely different from thermal conditions on Earth. Without an atmosphere, the only means of exchanging thermal energy is by thermal radiation and conduction. Certain parts of satellites have been calculated to follow 2 The term ‘torr’ is generally used instead of ‘mm Hg’ by international agreement of several vacuum societies. ‘Torr’ honours the name of Torricelli, who discovered atmospheric pressure in 1643. 22 2 Requirements for Spacecraft Materials
- 45. thermal excursions from approximately −160 to +180 °C. The actual temperature attained will differ from one space- craft to another, the major temperature effect arising from the spacecraft’s spin rate. Surfaces of non-spinning satellites exposed to direct solar radiation may be unable to dissipate thermal energy efficiently, and will reach higher tempera- tures than spinning satellites. Temperature variations will also depend upon the amount of albedo radiation and the amount of thermal radiation to space from the spacecraft. Both active and passive thermal control systems are employed on satellites in order to restrict the oscillating temperature extremes. The active systems have made use of thermostatically controlled heaters. Passive systems involve the surface absorptance/emittance, α/ε, properties of material finishes. Solar reflectors have low α/ε ratios, being generally white paints or clear anodized aluminium. Black paints and inorganic black anodized aluminium wave α/ε values of approximately 1. Solar absorbers have an α/ε value greater than 1, and these are generally polished metals since the emittance values of uncoated metals are very low (0.l). Examples are described in Sect. 5.5. The majority of metals do not sublimate at normal spacecraft temperatures. However, as can be seen in Table 2.5, cadmium and zinc must be excluded from use as they will readily sublimate and could condense in the form of thin conductive deposits on electrical insulators, or opa- que deposits on optical components which may be situated within the satellite or on its external surfaces. All cadmium, zinc, or tin-plated surfaces, such as the protective finishes on equipment or components including commercial connectors, must be avoided, because, as will be described in Chap. 7, they are known to grow single-crystal whiskers exceeding 2 cm length in vacuum. Extreme care must be taken to ensure that these metals are not used in the corrosion proofing of any of the spacecraft’s components. Magnesium parts could pose sublimation problems after long exposure to vacuum at temperatures greater than 125 °C, and experi- ments have shown (Frankel 1969) that magnesium sheets held at 230 °C and 1 × 10−7 mm Hg for only 168 h (one week) became severely pitted and dramatically decreased in static strength properties. Because of their relatively high strength-to-weight ratios, magnesium alloys are often employed as structural parts, but it is essential that these parts are finished with an adequate plating or chemical conversion coating which will prevent both the corrosion of the part before launch and the subsequent sublimation problem in orbit. Tin–lead alloys, such as those employed for soldering electrical components, have not been seen to sublimate under spacecraft environments as they are restricted for use in areas which are thermally controlled to a maximum of about 80 °C. Solder alloys are used for joining silver-plated molybdenum interconnector strips between solar cells on the solar arrays of spinning satellites. As such satellites rotate, the maximum temperature of the arrays does not degrade the soldered joints. The stationary communica- tion satellites will not be able to dissipate the absorbed thermal radiation on the solar arrays efficiently, so that welded interconnectors are necessary. The events of sublimation and evaporation cause the release of metal atoms which travel and are capable of recondensing on cooler surfaces. They are readily ionized and may be contributors to corona and arcing phenomena. These metallic ions may also cause the complete failure of a satellite mission by recondensing between slip rings, causing electrical short circuits, or recondensing on optical surfaces causing the loss of a specific wavelength transmission. When such ions recondense on the spacecraft’s highly reflective thermal control surfaces the thermal balance can be so Table 2.4 Characteristics of the space environment—order of magnitude only (Dauphin 1984) Altitude (km) Pressure (mm Hg or torr) Kinetic temperature (K) Gaseous density (particle cm−1 ) Composition Ultraviolet radiation Particle radiation (particles cm−2 s−1 ) Sea level 760 ±300 2.5 × 1019 78 % N2, 21 % O2, 1 % A Section of solar spectrum 0.3 – 30 10 – 4 × 1017 N2, O2, A Absorption zone – 200 10−6 ±1200 1010 N2, O, O2, O+ Full solar spectrum – 800 10−9 ±1300 106 O, He, O+ , H Full solar spectrum – 6500 10−13 – 103 H+ , H, He+ Full solar spectrum 104 protons 35 MeV 104 electrons 40 keV 22,000 10−13 – 101 –102 85 % H+ , 15 % He2+ Full solar spectrum 108 protons 5 MeV 108 electrons 40 keV 104 electrons 1.6 MeV 2.3 The Effect of a Space Environment 23
- 46. degraded that severe overheating will promote malfunctions. Cesium is an uncommon metal with a melting point of 28 °C and a boiling point of 671 °C but it has been proposed as a propellant for field effect electronic propulsion (FEEP) when very low thrust applications are required. Thrusters have been developed using this metal, but CNES evaluations have determined that cesium can easily contaminate spacecraft surfaces due to its low vapor pressure and re-evaporate from surfaces warmer than −30 °C to cause further contamination and react chemically with polymers or some oxides to Table 2.5 Sublimation of metals and semiconductors in high vacuum Sublimation rate 1000 Åa /year 10−3 cm/year (0.0004 in/year) 10−1 cm/year (0.04 in/year) Temperature °C °C °C Cadmium 40 80 120 Selenium 50 80 120 Zinc 70 132 180 Magnesium 110 170 240 Tellurium 130 180 220 Lithium 150 210 280 Antimony 210 270 300 Bismuth 240 320 400 Lead 270 330 430 Indium 400 500 610 Manganese 450 540 650 Silver 480 590 700 Tin 550 660 800 Aluminium 550 680 810 Beryllium 620 700 840 Copper 630 760 900 Gold 660 800 950 Germanium 660 800 950 Chromium 750 870 1000 Iron 770 900 1050 Silicon 790 920 1080 Nickel 800 940 1090 Palladium 810 940 1100 Cobalt 820 960 1100 Titanium 920 1070 1250 Vanadium 1020 1180 1350 Rhodium 1140 1330 1540 Platinum 1160 1340 1560 Boron 1230 1420 1640 Zirconium 1280 1500 1740 Iridium 1300 1500 1740 Molybdenum 1380 1630 1900 Carbon 1530 1680 1880 Tantalum 1780 2050 2300 Rhenium 1820 2050 2300 Tungsten 1880 2150 2500 a lÅ = 10−10 m Based on data by Jaffe and Rittenhouse, California Institute of Technology 24 2 Requirements for Spacecraft Materials
- 47. change their optical properties (Tondu 2011). The BepiColumbo mission dedicated to the study of Mercury is a challenging ESA project where temperatures may reach 300 °C at all external surfaces and metallic sublimation is of major concern (as is the outgassing of organic materials). A reference to Table 2.3 will immediately show that mate- rials containing zinc, cadmium (always prohibited in space applications) and lead are unsuitable candidates. Sublimation and cold welding of solar array drive mechanisms will require strict controls for the bearings, slip-rings and cables (Fink et al. 2009a, b). For additional examples of the effect of sublimation on spacecraft hardware, the reader may refer to Sect. 5.6. (c) Radiation and particle damage Organic materials, such as those used for electrical insula- tion, may be damaged by ionization due to protons and electrons from radiation belts, solar emissions, and cosmic rays. The Van Allen radiation belt is especially damaging to organic materials and even inorganic materials which make up optical lenses, ceramic insulators, and sensitive electronic components. The metallic materials of most Earth-orbiting spacecraft, as well as deep-space probes, are unaffected by small particle radiations, and are only slightly degraded by erosion from a cloud of meteoric dust which surrounds the Earth and other planets. Larger particle damage was, how- ever, a major problem for the Halley’s Comet fiyby mission —called Giotto—which encountered that comet in 1986. Special armour plating was developed to surround that spacecraft, which was impacted by rock debris travelling at hypervelocities of 10–70 km/s. These aspects are reviewed in Chap. 8. (d) Friction and wear One of the major material problem areas for advanced spacecraft is that of friction and wear of surfaces which must rub or slide over each other under conditions of temperature cycling and high vacuum. This may be encountered in the operation of hinges, gears, bearings, and electrical contacts used in a vast number of spacecraft mechanisms. During sliding under normal terrestrial conditions most contacting metallic surfaces are protected by a surface film of oxide, oil, grease, or other contaminant which will act as a ‘shear layer’ and prevent binding. Under vacuum conditions such contaminants outgas, and oxides, once disrupted or removed, are unable to reform. Also, the minute junction spots which carry the full load between contacting metallic surfaces will usually have vastly increased friction coeffi- cients, and probably high rates of wear, so that many metallic couples will tend to cold-weld. It is usually impracticable to enclose such moving parts within hermet- ically sealed containers, so special lubricants must be found which do not decompose or sublimate under vacuum. Films of low shear strength, such as molybdenum disulphide or vacuum deposited soft metals (e.g. gold, lead, or silver), are most efficient, particularly when designed to be situated between hard substrates which will support the load and keep the contact area small. The possibility of contacting metallic surfaces becoming cold welded to each other during, for instance, the operation of spacecraft mechanisms or loading of threaded fasteners under vacuum conditions depends on a number of factors; there will be a greater chance of cold welding if (a) the relative phase diagram indicates that the contacting metals or alloys form a solid solution with each other, (b) the metals are soft and have the same crystal structure, (c) contact surfaces are clean, or possess easily damaged/removable oxide films. Surface oxides are normally more brittle than metals and are therefore more likely to crack and expose underlying metal if it is too soft to provide a firm support under load, and (d) contact pressures are higher. The first step in order to minimize the possibility of cold welding is to select those metal combinations known to resist adhesive wear, as for instance those shown in Fig. 2.4a. The next step is to consider the surface finish. Clean metallic surfaces react with the surrounding atmosphere to form oxides, nitrides, or other compounds that are held together by either strong chemical bonds, or weak van der Waals forces. These surface films reduce the pos- sibility of metal–metal adhesion that would otherwise occur on intimate contact; they can be considered as naturally formed lubricating films. Under the space environment such films are unlikely to be self-healing if they become dis- placed. For this reason protective films of PTFE (Teflon), graphite, and molybdenum disulphide are frequently selec- ted to prevent wear and cold-welding. Suitable non-lubricated pairs of engineering alloys for sliding wear situations can be selected from engineering alloys having very different hardnesses—traditional refer- ence books may be consulted (Brandes 1983; Lansdown and Price 1986), or for very precise data, the recent work of Merstallinger et al. (2009) can be checked for evidence of fretting wear and cold welding (it is intended that this ref- erence work is maintained on the internet and updated as new pairs of engineering material pairs become tested). In general, steels can be coupled with either copper alloys or steels of a different alloy type and hardness. Copper alloys can be coupled with chromium plating, high-chrome steels, and tungsten steels. Austenitic stainless steels have a 2.3 The Effect of a Space Environment 25
- 48. great tendency to cold-weld to each other (see Fig. 2.4) as they are both soft and unable to form thick protective chromium oxides. Alternatively, with very hard substrates such as chromium plating, surface deformation is so small that the surface chromium oxide film is never disrupted (see Fig. 2.5). Materials suitable for spacecraft bearing applica- tions are titanium-carbide coated balls located in raceways fabricated from 440C or SAE 52100. A particularly good anti-wear surface, as for instance in gears, bushes, and piv- ots, is plasma-nitrided steel (Rowntree and Todd 1988). Thermal spraying is also a novel process for the coating of spacecraft subsystems enhancing wear resistance and as a thermal barrier—here thermal spraying can deposit both low and high melting point materials such as polymers and ceramics and metallic layers such as aluminium onto CFRP substrates such as antennae face-skins (Sturgeon and Dunn 2006; Saber-Samandari and Berndt 2010). Case histories related to wear are detailed in Sects. 5.2.7, 5.11, and 5.12. A large number of rules and design recommendations for the avoidance of wear and cold welding (for instance at mech- anism end stops, hold-down and release springs, sliding contacts, and ball bearings) have been listed in the form of a standard (Labruyère and Urmston 1995; Doyle and Hubbard 2010; ECSS-E-ST-32-08 2013). It should be noted that all dry lubricants wear, and will possess a finite life. However, improved wear characteristics can be achieved by carefully selecting the process of applying the dry lubricant. Ion-plated lead and sputtered molybdenum disulphide are now well proven, having low coefficients of friction and a long life. Burnished or spray-bonded molyb- denum disulphide has inferior friction properties. It is important to store the dry lubricated spacecraft mechanisms in a dry inert gas in order to prevent moisture pick-up, and as neither lead nor molybdenum disulphide perform well in air the number of operations in normal atmosphere should be restricted (Rowntree and Todd 1988). (e) Cryogenic temperatures All spacecraft structural metals will undergo changes in properties when cooled from normal ambient temperatures to temperatures in the ‘subzero’ range encountered during solar eclipse periods or when voyaging on deep-space mis- sions. This will be an important factor when liquid helium Fig. 2.4 a Depicts the theoretical possibility of clean metal surface pairs becoming cold-welded upon contact. Choice of a metal to resist adhesive wear with another specified metal. The “blacker” the circle, the better will be resistance to adhesive wear and cold welding under vacuum (Lansdown and Price 1986). Same-to-same metal contacts will result in solid solution, sticking and cold welding. b Demonstration of a case where “same metal” contacts have become cold welded (a form of solid state diffusion). By applying lead or silver coatings (e.g. as solid lubricants) such sticking will be avoided. The very thin oxide film on austenitic stainless steel is easily ruptured under the sliding conditions of torquing-up nuts and bolts. High vacuum operation has contributed to the complete seizure by cold-welding of this 316 alloy vacuum chamber support fixture. The 45 mm diameter bolt was initially cut to release the threaded portion which was later cross-sectioned. The pointer shows the main region of cold-welded asperities 26 2 Requirements for Spacecraft Materials
- 49. cryostats form a major part of a payload, as for instance has been designed for the Infrared Space Observatory, where sophisticated instruments are located in a 60 cm telescope cooled to 2 K. The greater changes involve the embrittle- ment of metal alloys, particularly carbon steels. Space vehicles must be fabricated from materials with high strength-to-weight ratios. They must also be required to retain high levels of fracture toughness at all service tem- peratures to ensure ‘fail safe’ lifetimes. In general, yield strengths, Young’s modulus, and tensile strengths increase as the exposed temperature is decreased. The effect of low-temperature exposure on ductility and toughness is, however, dependent on alloy composition, and for specific alloy data special handbooks should be consulted (Campbell 1980; Reed and Clark 1983). (f) Corrosion It should be emphasized that several effects of the space environments are beneficial to metallic materials. Before launch many criteria have to be set forth for the selection of spacecraft materials, so that failure resulting from corrosion and particularly stress-corrosion cracking will be prevented. With the exception of pressure vessels, plumbing lines, liquid fuel cells, and galvanic battery cells, these problems of cor- rosion are not evidenced in the vacuum environment of space. (g) Material fatigue The low and high cycle fatigue lives of parts fabricated from most steels, aluminium and titanium alloys are impressively extended under vacuum conditions—this is particularly welcome as many spacecraft parts will be subjected to extensive mechanical and thermal fatigue during their operational lives. An analysis of the results from extensive test programmes (Grinberg 1982) strongly indicates that the vacuum environment produces a change in the plastic strain intensity in the near-fatigue crack region of all alloys, and an increase in the plastic zone depth of ductile materials. This results in a decrease in crack propagation rate, due in part to the absence or a considerably reduced effectiveness of oxide or chemisorbed films on fresh crack surfaces. (h) Spacecraft charging In orbit or in deep space, spacecraft and space vehicles can develop an electric potential up to tens of thousands of volts relative to the ambient extraterrestrial plasma (the solar wind). These large potential differences (called ‘differential charging’) can also occur on the external surface of a launch vehicle. The main consequences of spacecraft differential charging are the phenomena of electrical discharge (‘coro- na’—which produces a damaging glow around conducting materials at high potential) and arcing (a luminous bridge formed by discharge between spacecraft electrical conduc- tors). Similar discharges may also be observed when high-voltage equipment, such as travelling wave tubes and electronic power supplies, operates on board for the trans- mission of signals from the spacecraft back to Earth. Many factors contribute to spacecraft charging, including the spacecraft configuration, its structural and surface materials, how correctly these materials are grounded, whether the craft is operating in sunlight or shadow, its altitude above Earth, and the flux density of high-energy solar particles or level of magnetic storm activity. Many possibilities exist to neu- tralize the spacecraft potential: where possible all intercon- necting parts, particularly at the surface, should be electrically grounded to ensure sufficient electrical conduc- tivity between interfaces. This will include solar call cover glasses and optical solar reflectors (see Sect. 5.5.4). Alter- native, new methods that reduce surface potentials (partic- ularly for scientific spacecraft designed to measure plasma and electric fields in the space environment) are active sys- tems that release a sufficient amount of charged particles across the external surfaces. These particles are ions, emitted Fig. 2.5 Example of 300 mm long Spacelab pallet trunnion— machined from Inconel 718 and hard chromium plated. The insert shows plating to be 10 μm thick and well bonded to the etched substrate 2.3 The Effect of a Space Environment 27
- 50. by field emission from a liquid metal source that can be indium. Ion release is usually for a short time, until the spacecraft potential is reduced, or reaches zero. (i) Spacecraft in hibernation The Rosetta spacecraft was manufactured by a European consortium during the 1998–2003 time period. It was laun- ched in March 2004 with the objective of making a ren- dezvous with the comet Churyumov-Gerasimenko, otherwise known as 67P. This craft achieved a new “first” in human history, by reaching its destination in August 2014 and began orbiting 100 km above the surface of the icy comet, taking images and then descending to a height of 30 km before detaching its lander, named Philae. Much of Rosetta’s 10 year journey was in hibernation mode. The journey covered 6.4 bn km through the Solar System (three times around the Earth, once around Mars, once close to Jupiter and five times around the Sun). Consequently this spacecraft’s hardware was subjected to most of the space environments compiled into (b) to (h) above. Close to the Sun the problem of overheating was solved by using radi- ators to dissipate heat into Space. Conversely, close to Jupiter, the hardware and experiments (20 in all) were kept warm by multi-layer insulation blankets and heaters located at strategic points such as fuel tanks, pipework and thrusters. The Philae lander was separated from Rosetta by means of a small pyro (explosive) cable cutter which activated the release of a large compressed spring. Much metallurgical work was conducted prior to launch to ensure that the Car- penter spring steel would not become embrittled at the very low (−160 °C) outer Solar System temperatures, or become cold welded to its mated structural surfaces, during its pas- sage close to the Sun. Philae’s landing gear was also mate- rially demanding due to the low temperatures encountered and low power budget (Thiel et al. 2003). The main com- ponents being harpoons to anchor to the comet’s surface: a copper beryllium projectile, pyrotechnic expansion system, cable magazine and a rewind system (AA 7075-T7351) driven by a brushless motor having plain bearings machined from MoS2-filled polyimide (Vespel SP3). Rosetta is supplied by power from two 14-m-long solar arrays having a total area of 64 m2 . The Si solar cells used are 200 µm thick, of low intensity, low temperature type, approximately 38 × 62 mm size. The cover glasses are 100 µm thick ceria doped micro-sheets. Four 10 Ah NiCd batteries store the power to supply the 28 V bus lines. Many other scientific spacecraft covering great distances use power from radioisotope thermoelectric generators (RTG). For instance, the New Horizons spacecraft, launched in January 2006 has been in hibernation for two thirds of its flight time, and will reach Pluto in mid-2015. On reaching the most outer bodies known to orbit the Sun, the RGA power system will be turned on and seven science instru- ments activated. 2.4 Materials for Space Launch Vehicles At present the only way that satellites, people, and cargo can be carried off from the Earth into the environment of space is by the use of rocket-propelled vehicles. Expendable launch vehicles (ELVs) are used only once. Many nations are involved with the construction and launch of ELVs. The most well known of the several hundred launch vehicles to have boosted spacecraft from Earth are listed in Table 2.6. The first European telecommunications satellite (OTS) was lost when its ELV exploded during launch, probably due to a defective solid rocket motor case. The failure review established that the steel case material had been incorrectly heat-treated. Parts of the exploded case were retrieved from the Atlantic Ocean by submarine. Metallographic evidence determined that the large cylindri- cal piece-parts had received an austenitizing time or tem- perature which was insufficient to solution treat the AISI 4130 (0.3C, 0.95Cr, 0.2Mo rem. Fe) steel. This was apparent from the presence of large-sized spherical carbides in the microstructure of the solid rocket motor case. The flight hardware case had not achieved peak hardness during sub- sequent quenching and normalizing. Incidentally, the in-line process control sample that had accompanied the flight case did have adequate mechanical and microstructural properties —this was because of the small mass of the test piece which responded well to the time-temperature profile. A replace- ment spacecraft, OTS 2, was launched successfully 8 months later in 1978 on Thor Delta number 141—this event can be seen in Fig. 2.2. The Delta ELV continues to be one of the most successful US launchers. A schematic diagram of the main features of a typical ELV is shown in Fig. 2.6. This shows the Titan rocket which was initially developed in the USA during the 1950s and continues to be launched today as a stretched version (Titan in and IV). This ELV, together with the Delta rocket, was complementary to the Space Shuttle fleet, particularly for the launch of heavy payloads. The fleet of Space Shuttles were retired in 2011 and NASA has selected two spacecraft to potentially replace the shuttles. It is intended that the US will take astronauts to the International Space Station (ISS) in 2017 by means of reusable capsules, the SpaceX Dragon and Boeing’s CST-100. Each capsule can be placed on single-use rockets, such as the Falcon 9 Heavy or Atlas 5 series, and they have been designed to carry up to seven astronauts at once. At the time of writing, only Russia is able to transport astronauts to and from the ISS by means of its 28 2 Requirements for Spacecraft Materials
- 51. Soyuz rocket. Supplies of equipment and food are trans- ported to the ISS using either the Russian Progress space- craft of the European Automatic Transfer Vehicle. Since its first flight in 2008, the ATV has played a vital role in ISS logistics serving as a cargo carrier, ‘space tug’ and storage facility. ATV will evolve into the European Service Module (ESV) designed to support the NASA Orion spacecraft. The deep-space missions envisaged by NASA will rely on the Orion spacecraft—it has been flown on a successful test flight aboard a Delta 4 Heavy booster launched from the Kennedy Space Centre at the end of 2014. Propellants for launch vehicles are regarded as materials. All rockets are propelled into the vacuum of space by utilizing only the liquid or solid materials on-board; at pre- sent, there is no possibility of using atmospheric oxygen. The term ‘propellant’ is used to denote the two chemical products, the ‘oxidizer’ and the ‘fuel’, contained inside conventional rockets. The fuel is burnt with the oxidizer in order to achieve the enormous amounts of energy needed for liftoff. The propellants can be in either liquid or solid state. Modern rockets are powered by ‘cryogenic’ propellants: liquids which at atmospheric pressure have boiling points below 0 °C. Examples are liquid oxygen (−183 °C) and liquid hydrogen (−253 °C). The most simple solid propellant is prepared from a mixture of nitrocellulose and nitroglycerine—this is the fuel. Table 2.6 Selected launch vehicles Country of origin Launch vehicle (latest known version) Typea Payload into GTOb (kg) China Long March 3 (1992) ELV 2500 Long March 3B (2007) ELV 11,500 (LEO) 5500 European Space Agency Ariane 4 (1990) ELV 2600 Ariane 5 (1996) ELV 6800 Ariane 5 (2014) ELV 10,500 Vega—Italy (2014) ELV 1500 (LEO) India Vehicle 3 (1979) ELV 40 (LEO) Polar Satellite Launch Vehicle (2014) ELV 3250 (LEO) Israel Shavit (1988) ELV 160 (LEO) Japan H-1 (1986) ELV 1100 H-IIB (2009) ELV 16,500 (LEO) 8000 USA Scout (1979) ELV 5400 Atlas 2 (1991) ELV 2700 Atlas 531 (2014) ELV 17,000 (LEO) Thor Delta (1992) ELV 2000 Delta 2 (2012) ELV 2500 Saturn V ELV 10,000 Titan III and IV (1989) ELV 5000 Falcon 9 (2014) ELV 13,150 (LEO) Space Shuttle (1990, retired 2011) AV 25,000 (LEO) Former USSR Vostok (1960) ELV 5000 (LEO) Proton (1968) Russia ELV 5500 Soyuz-2.1b (2014) Russia ELV 3000 Soyuz-2.1b (2014) Russia ELV 8500 (LEO) Soyuz-2.1v (2013) Russia ELV 2800 Zenit 3SL (2002) Ukraine ELV 6000 Buran Retired AV 30,000 (LEO) Key a ELV expendable launch vehicle; AV aerospace vehicle b Geosynchronous transfer orbit [flight performance to Low Earth Orbit (LEO) is usually more than twice this payload weight] 2.4 Materials for Space Launch Vehicles 29
- 52. The oxidizer is prepared separately from either ammonium nitrate, ammonium perchlorate, or potassium perchlorate. The fuel and oxidizer are made into powder form and then mixed with binder such as polyvinyl chloride or a poly- urethane. The resulting substance is poured into the solid rocket motor casing, where it sets hard. The case is usually made of steel, as discussed previously for the Thor Delta solid rocket motor, but composite materials based on gra- phite fibres are also used. More complex solid propellant chemistry is based on polybutaidene acrylonitrile (PBAN) and hydroxytelechelique polybutadiene (HTPB) propellant. The HTPB propellant for Ariane V solid boosters is, typi- cally, 68 % ammonium perchlorate, 14 % polybutadiene, and 18 % aluminium powder. Much progress has been made during the last decade in the development of improved solid propellants. One propellant is identified as GAP/Al/HNF [glycidyl azide polymer being the binder, aluminium powder and a new form of powerful oxidizer with the chemical composition of hydrazinium nitroformate (Schoyer 1996)]. Comparative tests were made between this new propellant and the best-performing HTPB-containing propellant. The findings included an increase in characteristic velocity of about 8 % and, importantly, an ecologically benign exhaust, free of chlorine, where the combustion products are nitrogen, water, carbon dioxide, nitrogen oxide, and aluminium oxide. Liquid propellant motors are usually fed from two tanks, one containing the liquid fuel (such as kerosene, liquid hydrogen, or hydrazine (N2H4)), the other containing the oxidizer (usually liquid oxygen or nitrogen tetroxide (N2O4)). Infrequently, fuming red nitric acid or nitrogen peroxide is used. The propellants are injected into the rocket motor’s combustion chamber, where ignition and combustion occur with a great release of thermal energy. The combustion gases are then forced through the exit nozzle of the motor. Here, the kinetic energy is absorbed by the nozzle as the gas velocity increases and then decreases through the ‘throat’ of the nozzle. A drawing of the Ariane IV launcher vehicle parts is shown in Fig. 2.7. The first two stages of this ELV use hydrazine and nitrogen tetroxide. The third stage is propelled by a cryogenic engine fed by liquid hydrogen and liquid oxygen. As illustrated in the figure, Ariane IV has the pos- sibility to incorporate either liquid or solid propellant booster motors to assist in the first stage liftoff. In contrast to Ariane, the Space Shuttle was, as its name implied, reusable and of extreme importance for its role in manned, near-Earth activities. The Shuttle’s primary propulsion consisted of a large external tank (47 m in length and 8.7 m in diameter) which contained compartments for the liquid oxygen/liquid hydrogen propellants—these were fed to the three main engines, and two strap-on booster motors which contained a solid composite propellant (polybutadiene, acrylic binder, and ammonium perchlorate oxidizer). Rocket structural materials are usually based on the Duralumin series of aluminium alloys (4 %Cu, 2 %Mn, rem. Al). These alloys have high strength-to-weight ratios and are detailed in further sections of this book. Concerning Ari- ane IV, much use has been made of the aluminium alloys AU4GN (AA2024), AZ5GU (AA7075), and AZ5G (AA7020). These are the French alloy designations with the US Aluminium Association designation in parentheses. Further cross-references to national alloy specifications are given in Appendix 6. Some locations for these alloys are indicated on Fig. 2.7. Unfortunately, several of their heat-treatment conditions are susceptible to stress corrosion cracking (SCC). A particular problem resulting from SCC is discussed in Sect. 4.5. All the major structural materials used for the construction of Ariane and its motors are identified in Fig. 2.7. The more modern rocket motor and booster bodies containing solid propellants are machined and welded from ‘maraging’ steels, which are based on iron with large amounts of nickel, cobalt, and molybdenum. These steels Fig. 2.6 View of the main structural parts of an ELV (based on a Titan III design) 30 2 Requirements for Spacecraft Materials
- 53. can be easily rolled and formed into complicated shapes, then welded—after this a suitable heat treatment is made which produces very hard and tough material properties. Liquid fuels and oxidizers are usually stored in pressure tanks which may be made of titanium or aluminium alloys, as shown in Fig. 2.8. These inner surfaces which make contact with the liquid fuels must be tested and found compatible and non-igniting with the liquid. For instance, the titanium alloy Ti6A14V is known to be compatible with hydrazine, and aluminium alloys compatible with liquid oxygen. Similarly, any pressure vessel liner materials based on organic materials need to be compatibility tested—this is more difficult, only one resin system, Torlon Al-10, being found to be compatible with liquid oxygen (Healy et al. 1995). The Zenit ELV of the former Soviet space agency was formally announced in 1989. The vehicle is a two-stage liquid oxygen and kerosene rocket which, like all CIS launchers, is assembled horizontally. Its launcher assembly, payload integration, and launch preparation phases have Fig. 2.7 Launch vehicle inboard profile with main structural materials indicated (Ariane IV, 42 LP) 2.4 Materials for Space Launch Vehicles 31
- 54. been described as ‘highly automated’ (Isakowitz 1995). General views of the Zenit assembly hall and its kerosene– LOX engine are seen in Figs. 2.9 and 2.10 respectively. This ELV has recently been proposed for launching satellites from the sea. The novel idea is to utilize a covered semi-submersible oil rig as the firing platform. This would be anchored in the Pacific, at the Equator, so making full advantage of the Earth’s maximum rotational velocity (about 1600 kph). The Zenit rocket, standing 62 m tall, would face eastward in a direction avoiding any inhabitable landfall. All other major world space launch sites are rather far north from the Equator with the exception of ESA’s facility at Kourou, French Guiana, which is less than 500 km north of the Equator. The uniquely flexible ‘Sea Launch’ site will undoubtedly require the implementation of a comprehensive corrosion protection scheme for all the associated spacecraft materials. The Ariane V development was initiated by the European Space Agency in 1985. It is designed for commercial mis- sions to launch satellites and cargo for the future space stations. The proposal for launch of the European Hermes as a reusable winged manned space vehicle atop the Ariane V was cancelled in the mid 90’s. However, it has flown several cargo vehicle [Automated Transfer Vehicle (ATV)] and is designed for a Crew Transfer Vehicle (CTV). An illustration of the Ariane V in a dual launch configuration is shown in Fig. 2.11. Ariane V has a length of 54 m, a gross mass of 710,000 kg, and a designed thrust at liftoff of 15.9 MN. The Vulcain engine powers the Ariane V main cryogenic stage. This engine consists of a gas generator cycle, in which tur- bopumps driven by a gas generator fed by propellants tapped from the main supply system feed fuel and oxidizer to the combustion chamber. Liquid oxygen (oxidizer) and liquid hydrogen (fuel) are sprayed into the combustion chamber. Because of the extremely high combustion temperature, reaching 3600 °C and with about 1600 °C at the internal wall of the chamber, it is necessary to cool the chamber. This is done by machining channels into the chamber wall—there are 360—and passing liquid hydrogen through them as the engine is fired. The chamber is fabricated from a wrought high-strength copper alloy (Narloy-Z) with an outer band of nickel. The combustion chambers of the Space Shuttle main engines were also made from Narloy-Z (Cu, 3Ag, 0.5Zr wt% with O2 of approximately 50 ppm) and, when correctly heat treated, this alloy was ideal for combustion chambers oper- ating from −252 to 540 °C but above this temperature the hot wall mechanical properties were degraded because of grain growth, grain boundary sliding and bulging that Fig. 2.8 A liquid propellant rocket engine Fig. 2.10 Installation of Zenit first stage kerosene—LO2 engine Fig. 2.9 General view of the Zenit assembly hall 32 2 Requirements for Spacecraft Materials
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- 56. principium individuationis, like a kaleidoscope, shows us in ever- shifting evanescent forms, there is an underlying unity, not only truly existing, but actually accessible to us; for lo! in tangible, objective form, it stands before our sight. Of these two mental attitudes, according as the one or the other is adopted, so the ϕιλία (Love) or the νεῑκος (Hatred) of Empedocles appears between man and man. If any one, who is animated by νεῑκος, could forcibly break in upon his most detested foe, and compel him to lay bare the inmost recesses of his heart; to his surprise, he would find again in the latter his very self. For just as in dreams, all the persons that appear to us are but the masked images of ourselves; so in the dream of our waking life, it is our own being which looks on us from out our neighbours' eyes,—though this is not equally easy to discern. Nevertheless, tat tvam asi. The preponderance of either mode of viewing life not only determines single acts; it shapes a man's whole nature and temperament. Hence the radical difference of mental habit between the good character and the bad. The latter feels everywhere that a thick wall of partition hedges him off from all others. For him the world is an absolute non-ego, and his relation to it an essentially hostile one; consequently, the key-note of his disposition is hatred, suspicion, envy, and pleasure in seeing distress. The good character, on the other hand, lives in an external world homogeneous with his own being; the rest of mankind is not in his eyes a non-ego; he thinks of it rather as myself once more. He therefore stands on an essentially amicable footing with every one: he is conscious of being, in his inmost nature, akin to the whole human race,[9] takes direct interest in their weal and woe, and confidently assumes in their case the same interest in him. This is the source of his deep inward peace, and of that happy, calm, contented manner, which goes out on those around him, and is as the presence of a good diffused. Whereas the bad character in time of trouble has no trust in the help of his fellow-creatures. If he invokes aid, he does so without confidence: obtained, he feels no real gratitude for it; because he can hardly discern therein anything but the effect of others' folly. For
- 57. he is simply incapable of recognising his own self in some one else; and this, even after it has furnished the most incontestible signs of existence in that other person: on which fact the repulsive nature of all unthankfulness in reality depends. The moral isolation, which thus naturally and inevitably encompasses the bad man, is often the cause of his becoming the victim of despair. The good man, on the contrary, will appeal to his neighbours for assistance, with an assurance equal to the consciousness he has of being ready himself to help them. As I have said: to the one type, humanity is a non- ego; to the other, myself once more. The magnanimous character, who forgives his enemy, and returns good for evil, rises to the sublime, and receives the highest meed of praise; because he recognises his real self even there where it is most conspicuously disowned. Every purely beneficent act all help entirely and genuinely unselfish, being, as such, exclusively inspired by another's distress, is, in fact, if we probe the matter to the bottom, a dark enigma, a piece of mysticism put into practice; inasmuch as it springs out of, and finds its only true explanation in, the same higher knowledge that constitutes the essence of whatever is mystical. For how, otherwise than metaphysically, are we to account for even the smallest offering of alms made with absolutely no other object than that of lessening the want which afflicts a fellow-creature? Such an act is only conceivable, only possible, in so far as the giver knows that it is his very self which stands before him, clad in the garments of suffering; in other words, so far as he recognises the essential part of his own being, under a form not his own.[10] It now becomes apparent, why in the foregoing part I have called Compassion the great mystery of Ethics. He, who goes to meet death for his fatherland, has freed himself from the illusion which limits a man's existence to his own person. Such a one has broken the fetters of the principium individuationis. In his widened, enlightened nature he embraces all his countrymen, and in them lives on and on. Nay, he reaches forward to, and
- 58. merges himself in the generations yet unborn, for whom he works; and he regards death as a wink of the eyelids, so momentary that it does not interrupt the sight. We may here sum up the characteristics of the two human types above indicated. To the Egoist all other people are uniformly and intrinsically strangers. In point of fact, he considers nothing to be truly real, except his own person, and regards the rest of mankind practically as troops of phantoms, to whom he assigns merely a relative existence, so far as they may be instruments to serve, or barriers to obstruct, his purposes; the result being an immeasurable difference, a vast gulf between his ego on the one side, and the non-ego on the other. In a word, he lives exclusively centred in his own individuality, and on his death-day he sees all reality, indeed the whole world, coming to an end along with himself.[11] Whereas the Altruist discerns in all other persons, nay, in every living thing, his own entity, and feels therefore that his being is commingled, is identical with the being of whatever is alive. By death he loses only a small part of himself. Patting off the narrow limitations of the individual, he passes into the larger life of all mankind, in whom he always recognised, and, recognising, loved, his very self; and the illusion of Time and Space, which separated his consciousness from that of others, vanishes. These two opposite modes of viewing the world are probably the chief, though not indeed the sole cause of the difference we find between very good and exceptionally bad men, as to the manner in which they meet their last hour. In all ages Truth, poor thing, has been put to shame for being paradoxical; and yet it is not her fault. She cannot assume the form of Error seated on his throne of world-wide sovereignty. So then, with a sigh, she looks up to her tutelary god, Time, who nods assurance to her of future victory and glory, but whose wings beat the air so slowly with their mighty strokes, that the individual perishes or ever the day of triumph be come. Hence I, too, am perfectly aware of the paradox which this metaphysical explanation of the ultimate ethical phaenomenon must present to Western minds, accustomed, as they are, to very different methods of
- 59. providing Morals with a basis. Nevertheless, I cannot offer violence to the truth. All that is possible for me to do, out of consideration for European blindness, is to assert once more, and demonstrate by actual quotation, that the Metaphysics of Ethics, which I have here suggested, was thousands of years ago the fundamental principle of Indian wisdom. And to this wisdom I point back, as Copernicus did to the Pythagorean cosmic system, which was suppressed by Aristotle and Ptolemaeus. In the Bhagavadgîtâ (Lectio XIII.; 27, 28), according to A. W. von Schlegel's translation, we find the following passage: Eundem in omnibus animantibus consistentem summum dominum, istis pereuntibus kaud pereuntem qui cernit, is vere cernit. Eundem vero cernens ubique praesentem dominum, non violat semet ipsum sua ipsius culpa: exinde pergit ad summum iter.[12] With these hints towards the elaboration of a metaphysical basis for Ethics I must close, although an important step still remains to be taken. The latter would presuppose a further advance in Moral Science itself; and this can hardly be made, because in the West the highest aim of Ethics is reached in the theory of justice and virtue. What lies beyond is unknown, or at any rate ignored. The omission, therefore, is unavoidable; and the reader need feel no surprise, if the above slight outline of the Metaphysics of Ethics does not bring into view—even remotely—the corner-stone of the whole metaphysical edifice, nor reveal the connection of all the parts composing the Divina Commedia. Such a presentment, moreover, is involved neither in the question set, nor in my own plan. A man cannot say everything in one day, and should not answer more than he is asked. He who tries to promote human knowledge and insight is destined to always encounter the opposition of his age, which is like the dead weight of some mass that has to be dragged along: there on the ground it lies, a huge inert deformity, defying all efforts to quicken its shape with new life. But such a one must take comfort from the certainty that, although prejudices beset his path, yet the truth is with him. And Truth does but wait for her ally, Time, to join her;
- 60. once he is at her side, she is perfectly sure of victory, which, if to- day delayed, will be won to-morrow. [1] The conception of the Good, in its purity, is an ultimate one, an absolute Idea, whose substance loses itself in infinity.—(Bouterweek: Praktische Aphorismen, p. 54.) It is obvious that this writer would like to transform the familiar, nay, trivial conception Good into a sort of Διἴπετής, to be set up as an idol in his temple. Διἴπετής lit., fallen from Zeus; and so heaven-sent, a thing of divine origin. Cf. Horn., Il.. XVI, 174; Od.. IV. 477. Eur., Bacch., 1268.—(Translator.) [2]The genuineness of the O u p n e k ' h a t has been disputed on the ground of certain marginal glosses which were added by Mohammedan copyists, and then interpolated in the text, it has, however, been fully established by the Sanskrit scholar, F.H.H. Windischmann (junior) in his Sancara, sive de Theologumenis Vedanticorum, 1833, p. xix; and also by Bochinger in his book De la Vie Contemplative chez les Indous, 1831, p. 12. The reader though ignorant of Sanskrit, may yet convince himself that Anquetil Duperron's word for word Latin translation of the Persian version of the U p a n i s h a d s made by the martyr of this creed, the Sultan D â r â - S h u k o h, is based on a thorough and exact knowledge of the language. He has only to compare it with recent translations of some of the U p a n i s h a d s by Rammohun Boy, by Poley, and especially with that of Colebrooke, as also with Röer's, (the latest). These writers are obviously groping in obscurity, and driven to make shift with hazy conjectures, so that without doubt their work is much less accurate. More will be found on this subject in Vol. II. of the Parerga, chap. 16, § 184. [V. The Upanishads, translated by Max Müller, in The Sacred Books of the East, Vols. I. and XV. Cf. also Max Müller, The Science of Language, Vol. I., p. 171. Now that an adequate translation of the original exists, the O u p n e k ' h a t has only an historical interest. The value which Schopenhauer attached to the U p a n i s h a d s is very clearly expressed also in the Welt als Wille und Vorstellung, Preface to the first Edition; and in the Parerga, II., chap, xvi., § 184.—(Translator.)] [3] For the S û f i, more correctly *Sūfīy a sect which appeared already in the first century of the H i j r a h, the reader is referred to: Tholuck's Blüthensammlung aus der Morgenländischen Mystik (Berlin, 1825); Tholuck's Sûfismus, sive Theosophia Persarum Pantheistica (Berlin, 1821); Kremer's Geschichte der Herrschenden Ideen des Islâms (Leipzig, 1868); Palmer's Oriental Mysticism (London, 1867); Gobineau's Les Religions et les Philosophies dans l'Asie Centrale (2nd edit. Paris, 1866); A Dictionary of Islâm, by T. P. Hughes (London, 1885), p. 608 sqq.— (Translator.)
- 61. [4] This is too well-known to need verification by references. The Cantico del Sole by St. Francis of Assisi sounds almost like a passage from the U p a n i s h a d s or the B h a g a v a d g î t â.—(Translator.) [5] On peut assez longtemps, chez notre espèce, Fermer la porte à la Raison. Mais, dès qu'elle entre avec adresse, Elle reste dans la maison, Et bientôt elle en est maîtresse. —(Voltaire.) (We men may, doubtless, all our lives To Reason bar the door. But if to enter she contrives, The house she leaves no more, And soon as mistress there presides.) [6] Τὸ ἔν= the eternal Reality outside Time and Space Tὸ πᾱν = the phaenomenal universe.—(Translator.) [7] Mâyâ is the delusive reflection of the true eternal Entity.—(Translator.) [8] This expression is used in the Brahmanical philosophy to denote the relation between the world-fiction as a whole and its individualised parts. V. A. E. Gough, Philosophy of the Upanishads, 1882.—(Translator.) [9] Homo sum: humani nil a me alienum puto. Terence, Heaut., I. 1, 25.— (Translator.) [10] It is probable that many, perhaps, most cases of truly disinterested Compassion—when they really occur—are due not to any conscious knowledge of this sort, but to an unconscious impulse springing from the ultimate unity of all living things, and acting, so to say, automatically.—(Translator.) [11] Cf. Richard Wagner: Jesus von Nazareth; pp. 79-90.—(Translator.) [12] That man is endowed with true insight who sees that the same ruling power is inherent in all things, and that when these perish, it perishes not. For if he discerns the same ruling power everywhere present, he does not degrade himself by his own fault: thence he passes to the highest path.—For the Bhagavadgîtâ the reader is referred to Vol. VIII. of The Sacred Books of the East (Oxford: Clarendon Press), where (p. 105) this passage is translated as follows:—He sees (truly) who sees the supreme lord abiding alike in all entities, and not destroyed though they are destroyed. For he who sees the lord abiding everywhere alike, does not destroy himself[*] by himself, and then reaches the highest goal. [*]Not to have true knowledge, is equivalent to self-destruction.
- 62. Cf. Fauche: Le Mahā-bhārata: Paris, 1867; Vol. VII., p. 128:— Celui-là possède une vue nette des choses, qui voit ce principe souverain en tous les êtres d'une manière égale, et leur survivre, quand ils périssent. Il ne se fait aucun tort à soi-même par cette vue d'un principe qui subsiste également partout: puis, après cette vie, il entre dans la voie supérieure. The obscurity of Schlegel's Latin in the second sentence is sufficiently removed by these more recent translations.—(Translator.) JUDICIUM REGIAE DANICAE SCIENTIARUM SOCIETATIS. Quaestionem anno 1837 propositam, utrum philosophiae moralis fons et fundamentum in idea moralitatis, quae immediate conscientia contineatur, et ceteris notionibus fundamentalibus, quae ex ilia prodeant, explicandis quaerenda sint, an in alio cognoscendi principio, unus tantum scriptor explicare conatus est, cujus commentationem, germanico sermone compositam, et his verbis notatam: MORAL PREDIGEN IST LEICHT, MORAL BEGRÜNDEN IST SCHWER, praemio dignam judicare nequivimus. Omisso enim eo, quod potissimum postulabatur, hoc expeti putavit, ut principium aliquod ethicae conderetur, itaqae eam partem commentationis suae, in qua principii ethicae a se propositi et metaphysicae suae nexum exponit, appendices loco habuit, in qua plus quam postulatum esset praestaret, quum tamen ipsum thema ejusmodi disputationem flagitaret, in qua vel praecipuo loco metaphysicae et ethicae nexus consideraretur. Quod autem scriptor in sympathia fundamentum ethicae constituere conatus est, neque ipsa disserendi forma nobis satisfecit, neque reapse, hoc fundamentum sufficere, evicit; quin ipse contra esse confiteri coactus est. Neque reticendum videtur, plures recentioris aetatis summos philosophos tam indecenter commemorari, ut justam et gravem offensionem habeat.
- 63. JUDGMENT OF THE DANISH ROYAL SOCIETY OF SCIENCES. In 1837 the following question was set as subject for a Prize Essay: Is the fountain and basis of Morals to be sought for in an idea of morality which lies directly in the consciousness (or conscience), and in the analysis of the other leading ethical conceptions which arise from it? Or is it to be found in some other source of knowledge? There was only one competitor; but his dissertation, written in German, and bearing the motto: To preach Morality is easy, to found it is difficult[1] we cannot adjudge worthy of the Prize. He has omitted to deal with the essential part of the question, apparently thinking that he was asked to establish some fundamental principle of Ethics. Consequently, that part of the treatise, which explains how the moral basis he proposes is related to his system of metaphysics, we find relegated to an appendix, as an opus supererogationis, although it was precisely the connection between Metaphysics and Ethics that our question required to be put in the first and foremost place. The writer attempts to show that compassion is the ultimate source of morality; but neither does his mode of discussion appear satisfactory to us, nor has he, in point of fact, succeeded in proving that such a foundation is adequate. Indeed he himself is obliged to admit that it is not.[2] Lastly, the Society cannot pass over in silence the fact that he mentions several recent philosophers of the highest standing in an unseemly manner, such as to justly occasion serions offence. [1] The Academy has been good enough to insert the second is on its own account, by way of proving the truth of Longinus' theory (V. De Sublimitate: chap. 39, ad fin.), that the addition or subtraction of a single syllable is sufficient to destroy the whole force of a sentence. (P. Longinus: De Sublimitate Libellus; edit. Joannes Vablen, Bonnae, 1887.)—(Translator) [2] I suppose this is the meaning of contra esse confiteri.— (Translator.)
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