Personal observations on the reliability of the Shuttle
by R. P. Feynman, 1986.
Introduction
It appears that there are enormous differences of opinion as to the
probability of a failure with loss of vehicle and of human life. The
estimates range from roughly 1 in 100 to 1 in 100,000. The higher
figures come from the working engineers, and the very low figures from
management. What are the causes and consequences of this lack of
agreement? Since 1 part in 100,000 would imply that one could put a
Shuttle up each day for 300 years expecting to lose only one, we could
properly ask "What is the cause of management's fantastic faith in the
machinery?"
We have also found that certification criteria used in Flight
Readiness Reviews often develop a gradually decreasing strictness. The
argument that the same risk was flown before without failure is often
accepted as an argument for the safety of accepting it again. Because
of this, obvious weaknesses are accepted again and again, sometimes
without a sufficiently serious attempt to remedy them, or to delay a
flight because of their continued presence.
There are several sources of information. There are published criteria
for certification, including a history of modifications in the form of
waivers and deviations. In addition, the records of the Flight
Readiness Reviews for each flight document the arguments used to
accept the risks of the flight. Information was obtained from the
direct testimony and the reports of the range safety officer, Louis
J. Ullian, with respect to the history of success of solid fuel
rockets. There was a further study by him (as chairman of the launch
abort safety panel (LASP)) in an attempt to determine the risks
involved in possible accidents leading to radioactive contamination
from attempting to fly a plutonium power supply (RTG) for future
planetary missions. The NASA study of the same question is also
available. For the History of the Space Shuttle Main Engines,
interviews with management and engineers at Marshall, and informal
interviews with engineers at Rocketdyne, were made. An independent
(Cal Tech) mechanical engineer who consulted for NASA about engines
was also interviewed informally. A visit to Johnson was made to gather
information on the reliability of the avionics (computers, sensors,
and effectors). Finally there is a report "A Review of Certification
Practices, Potentially Applicable to Man-rated Reusable Rocket
Engines," prepared at the Jet Propulsion Laboratory by N. Moore, et
al., in February, 1986, for NASA Headquarters, Office of Space
Flight. It deals with the methods used by the FAA and the military to
certify their gas turbine and rocket engines. These authors were also
interviewed informally.
Solid Rockets (SRB)
An estimate of the reliability of solid rockets was made by the range
safety officer, by studying the experience of all previous rocket
flights. Out of a total of nearly 2,900 flights, 121 failed (1 in
25). This includes, however, what may be called, early errors, rockets
flown for the first few times in which design errors are discovered
and fixed. A more reasonable figure for the mature rockets might be 1
in 50. With special care in the selection of parts and in inspection,
a figure of below 1 in 100 might be achieved but 1 in 1,000 is
probably not attainable with today's technology. (Since there are two
rockets on the Shuttle, these rocket failure rates must be doubled to
get Shuttle failure rates from Solid Rocket Booster failure.)
NASA officials argue that the figure is much lower. They point out
that these figures are for unmanned rockets but since the Shuttle is a
manned vehicle "the probability of mission success is necessarily very
close to 1.0." It is not very clear what this phrase means. Does it
mean it is close to 1 or that it ought to be close to 1? They go on to
explain "Historically this extremely high degree of mission success
has given rise to a difference in philosophy between manned space
flight programs and unmanned programs; i.e., numerical probability
usage versus engineering judgment." (These quotations are from "Space
Shuttle Data for Planetary Mission RTG Safety Analysis," Pages 3-1,
3-1, February 15, 1985, NASA, JSC.) It is true that if the probability
of failure was as low as 1 in 100,000 it would take an inordinate
number of tests to determine it ( you would get nothing but a string
of perfect flights from which no precise figure, other than that the
probability is likely less than the number of such flights in the
string so far). But, if the real probability is not so small, flights
would show troubles, near failures, and possible actual failures with
a reasonable number of trials. and standard statistical methods could
give a reasonable estimate. In fact, previous NASA experience had
shown, on occasion, just such difficulties, near accidents, and
accidents, all giving warning that the probability of flight failure
was not so very small. The inconsistency of the argument not to
determine reliability through historical experience, as the range
safety officer did, is that NASA also appeals to history, beginning
"Historically this high degree of mission success..."
Finally, if we are to replace standard numerical probability usage
with engineering judgment, why do we find such an enormous disparity
between the management estimate and the judgment of the engineers? It
would appear that, for whatever purpose, be it for internal or
external consumption, the management of NASA exaggerates the
reliability of its product, to the point of fantasy.
The history of the certification and Flight Readiness Reviews will not
be repeated here. (See other part of Commission reports.) The
phenomenon of accepting for flight, seals that had shown erosion and
blow-by in previous flights, is very clear. The Challenger flight is
an excellent example. There are several references to flights that had
gone before. The acceptance and success of these flights is taken as
evidence of safety. But erosion and blow-by are not what the design
expected. They are warnings that something is wrong. The equipment is
not operating as expected, and therefore there is a danger that it can
operate with even wider deviations in this unexpected and not
thoroughly understood way. The fact that this danger did not lead to a
catastrophe before is no guarantee that it will not the next time,
unless it is completely understood. When playing Russian roulette the
fact that the first shot got off safely is little comfort for the
next. The origin and consequences of the erosion and blow-by were not
understood. They did not occur equally on all flights and all joints;
sometimes more, and sometimes less. Why not sometime, when whatever
conditions determined it were right, still more leading to
catastrophe?
In spite of these variations from case to case, officials behaved as
if they understood it, giving apparently logical arguments to each
other often depending on the "success" of previous flights. For
example. in determining if flight 51-L was safe to fly in the face of
ring erosion in flight 51-C, it was noted that the erosion depth was
only one-third of the radius. It had been noted in an experiment
cutting the ring that cutting it as deep as one radius was necessary
before the ring failed. Instead of being very concerned that
variations of poorly understood conditions might reasonably create a
deeper erosion this time, it was asserted, there was "a safety factor
of three." This is a strange use of the engineer's term ,"safety
factor." If a bridge is built to withstand a certain load without the
beams permanently deforming, cracking, or breaking, it may be designed
for the materials used to actually stand up under three times the
load. This "safety factor" is to allow for uncertain excesses of load,
or unknown extra loads, or weaknesses in the material that might have
unexpected flaws, etc. If now the expected load comes on to the new
bridge and a crack appears in a beam, this is a failure of the
design. There was no safety factor at all; even though the bridge did
not actually collapse because the crack went only one-third of the way
through the beam. The O-rings of the Solid Rocket Boosters were not
designed to erode. Erosion was a clue that something was wrong.
Erosion was not something from which safety can be inferred.
There was no way, without full understanding, that one could have
confidence that conditions the next time might not produce erosion
three times more severe than the time before. Nevertheless, officials
fooled themselves into thinking they had such understanding and
confidence, in spite of the peculiar variations from case to case. A
mathematical model was made to calculate erosion. This was a model
based not on physical understanding but on empirical curve fitting. To
be more detailed, it was supposed a stream of hot gas impinged on the
O-ring material, and the heat was determined at the point of
stagnation (so far, with reasonable physical, thermodynamic laws). But
to determine how much rubber eroded it was assumed this depended only
on this heat by a formula suggested by data on a similar material. A
logarithmic plot suggested a straight line, so it was supposed that
the erosion varied as the .58 power of the heat, the .58 being
determined by a nearest fit. At any rate, adjusting some other
numbers, it was determined that the model agreed with the erosion (to
depth of one-third the radius of the ring). There is nothing much so
wrong with this as believing the answer! Uncertainties appear
everywhere. How strong the gas stream might be was unpredictable, it
depended on holes formed in the putty. Blow-by showed that the ring
might fail even though not, or only partially eroded through. The
empirical formula was known to be uncertain, for it did not go
directly through the very data points by which it was
determined. There were a cloud of points some twice above, and some
twice below the fitted curve, so erosions twice predicted were
reasonable from that cause alone. Similar uncertainties surrounded the
other constants in the formula, etc., etc. When using a mathematical
model careful attention must be given to uncertainties in the model.
Liquid Fuel Engine (SSME)
During the flight of 51-L the three Space Shuttle Main Engines all
worked perfectly, even, at the last moment, beginning to shut down the
engines as the fuel supply began to fail. The question arises,
however, as to whether, had it failed, and we were to investigate it
in as much detail as we did the Solid Rocket Booster, we would find a
similar lack of attention to faults and a deteriorating
reliability. In other words, were the organization weaknesses that
contributed to the accident confined to the Solid Rocket Booster
sector or were they a more general characteristic of NASA? To that end
the Space Shuttle Main Engines and the avionics were both
investigated. No similar study of the Orbiter, or the External Tank
were made.
The engine is a much more complicated structure than the Solid
Rocket Booster, and a great deal more detailed engineering goes into
it. Generally, the engineering seems to be of high quality and
apparently considerable attention is paid to deficiencies and faults
found in operation.
The usual way that such engines are designed (for military or
civilian aircraft) may be called the component system, or bottom-up
design. First it is necessary to thoroughly understand the properties
and limitations of the materials to be used (for turbine blades, for
example), and tests are begun in experimental rigs to determine
those. With this knowledge larger component parts (such as bearings)
are designed and tested individually. As deficiencies and design
errors are noted they are corrected and verified with further
testing. Since one tests only parts at a time these tests and
modifications are not overly expensive. Finally one works up to the
final design of the entire engine, to the necessary
specifications. There is a good chance, by this time that the engine
will generally succeed, or that any failures are easily isolated and
analyzed because the failure modes, limitations of materials, etc.,
are so well understood. There is a very good chance that the
modifications to the engine to get around the final difficulties are
not very hard to make, for most of the serious problems have already
been discovered and dealt with in the earlier, less expensive, stages
of the process.
The Space Shuttle Main Engine was handled in a different manner,
top down, we might say. The engine was designed and put together all
at once with relatively little detailed preliminary study of the
material and components. Then when troubles are found in the
bearings, turbine blades, coolant pipes, etc., it is more expensive
and difficult to discover the causes and make changes. For example,
cracks have been found in the turbine blades of the high pressure
oxygen turbopump. Are they caused by flaws in the material, the effect
of the oxygen atmosphere on the properties of the material, the
thermal stresses of startup or shutdown, the vibration and stresses of
steady running, or mainly at some resonance at certain speeds, etc.?
How long can we run from crack initiation to crack failure, and how
does this depend on power level? Using the completed engine as a test
bed to resolve such questions is extremely expensive. One does not
wish to lose an entire engine in order to find out where and how
failure occurs. Yet, an accurate knowledge of this information is
essential to acquire a confidence in the engine reliability in use.
Without detailed understanding, confidence can not be attained.
A further disadvantage of the top-down method is that, if an
understanding of a fault is obtained, a simple fix, such as a new
shape for the turbine housing, may be impossible to implement without
a redesign of the entire engine.
The Space Shuttle Main Engine is a very remarkable machine. It has
a greater ratio of thrust to weight than any previous engine. It is
built at the edge of, or outside of, previous engineering
experience. Therefore, as expected, many different kinds of flaws and
difficulties have turned up. Because, unfortunately, it was built in
the top-down manner, they are difficult to find and fix. The design
aim of a lifetime of 55 missions equivalent firings (27,000 seconds of
operation, either in a mission of 500 seconds, or on a test stand) has
not been obtained. The engine now requires very frequent maintenance
and replacement of important parts, such as turbopumps, bearings,
sheet metal housings, etc. The high-pressure fuel turbopump had to be
replaced every three or four mission equivalents (although that may
have been fixed, now) and the high pressure oxygen turbopump every
five or six. This is at most ten percent of the original
specification. But our main concern here is the determination of
reliability.
In a total of about 250,000 seconds of operation, the engines have
failed seriously perhaps 16 times. Engineering pays close attention to
these failings and tries to remedy them as quickly as possible. This
it does by test studies on special rigs experimentally designed for
the flaws in question, by careful inspection of the engine for
suggestive clues (like cracks), and by considerable study and
analysis. In this way, in spite of the difficulties of top-down
design, through hard work, many of the problems have apparently been
solved.
A list of some of the problems follows. Those followed by an
asterisk (*) are probably solved:
1.Turbine blade cracks in high pressure fuel turbopumps (HPFTP). (May have been solved.)
2.Turbine blade cracks in high pressure oxygen turbopumps (HPOTP).
3.Augmented Spark Igniter (ASI) line rupture.*
4.Purge check valve failure.*
5.ASI chamber erosion.*
6.HPFTP turbine sheet metal cracking.
7.HPFTP coolant liner failure.*
8.Main combustion chamber outlet elbow failure.*
9.Main combustion chamber inlet elbow weld offset.*
10.HPOTP subsynchronous whirl.*
11.Flight acceleration safety cutoff system (partial failure in a redundant system).*
12.Bearing spalling (partially solved).
13.A vibration at 4,000 Hertz making some engines inoperable, etc.
Many of these solved problems are the early difficulties of a new
design, for 13 of them occurred in the first 125,000 seconds and only
three in the second 125,000 seconds. Naturally, one can never be sure
that all the bugs are out, and, for some, the fix may not have
addressed the true cause. Thus, it is not unreasonable to guess there
may be at least one surprise in the next 250,000 seconds, a
probability of 1/500 per engine per mission. On a mission there are
three engines, but some accidents would possibly be contained, and
only affect one engine. The system can abort with only two
engines. Therefore let us say that the unknown suprises do not, even
of themselves, permit us to guess that the probability of mission
failure do to the Space Shuttle Main Engine is less than 1/500. To
this we must add the chance of failure from known, but as yet
unsolved, problems (those without the asterisk in the list
above). These we discuss below. (Engineers at Rocketdyne, the
manufacturer, estimate the total probability as 1/10,000. Engineers at
marshal estimate it as 1/300, while NASA management, to whom these
engineers report, claims it is 1/100,000. An independent engineer
consulting for NASA thought 1 or 2 per 100 a reasonable estimate.)
The history of the certification principles for these engines is
confusing and difficult to explain. Initially the rule seems to have
been that two sample engines must each have had twice the time
operating without failure as the operating time of the engine to be
certified (rule of 2x). At least that is the FAA practice, and NASA
seems to have adopted it, originally expecting the certified time to
be 10 missions (hence 20 missions for each sample). Obviously the best
engines to use for comparison would be those of greatest total (flight
plus test) operating time -- the so-called "fleet leaders." But what
if a third sample and several others fail in a short time? Surely we
will not be safe because two were unusual in lasting longer. The short
time might be more representative of the real possibilities, and in
the spirit of the safety factor of 2, we should only operate at half
the time of the short-lived samples.
The slow shift toward decreasing safety factor can be seen in many
examples. We take that of the HPFTP turbine blades. First of all the
idea of testing an entire engine was abandoned. Each engine number has
had many important parts (like the turbopumps themselves) replaced at
frequent intervals, so that the rule must be shifted from engines to
components. We accept an HPFTP for a certification time if two samples
have each run successfully for twice that time (and of course, as a
practical matter, no longer insisting that this time be as large as 10
missions). But what is "successfully?" The FAA calls a turbine blade
crack a failure, in order, in practice, to really provide a safety
factor greater than 2. There is some time that an engine can run
between the time a crack originally starts until the time it has grown
large enough to fracture. (The FAA is contemplating new rules that
take this extra safety time into account, but only if it is very
carefully analyzed through known models within a known range of
experience and with materials thoroughly tested. None of these
conditions apply to the Space Shuttle Main Engine.
Cracks were found in many second stage HPFTP turbine blades. In one
case three were found after 1,900 seconds, while in another they were
not found after 4,200 seconds, although usually these longer runs
showed cracks. To follow this story further we shall have to realize
that the stress depends a great deal on the power level. The
Challenger flight was to be at, and previous flights had been at, a
power level called 104% of rated power level during most of the time
the engines were operating. Judging from some material data it is
supposed that at the level 104% of rated power level, the time to
crack is about twice that at 109% or full power level (FPL). Future
flights were to be at this level because of heavier payloads, and many
tests were made at this level. Therefore dividing time at 104% by 2,
we obtain units called equivalent full power level (EFPL). (Obviously,
some uncertainty is introduced by that, but it has not been studied.)
The earliest cracks mentioned above occurred at 1,375 EFPL.
Now the certification rule becomes "limit all second stage blades
to a maximum of 1,375 seconds EFPL." If one objects that the safety
factor of 2 is lost it is pointed out that the one turbine ran for
3,800 seconds EFPL without cracks, and half of this is 1,900 so we are
being more conservative. We have fooled ourselves in three ways. First
we have only one sample, and it is not the fleet leader, for the other
two samples of 3,800 or more seconds had 17 cracked blades between
them. (There are 59 blades in the engine.) Next we have abandoned the
2x rule and substituted equal time. And finally, 1,375 is where we did
see a crack. We can say that no crack had been found below 1,375, but
the last time we looked and saw no cracks was 1,100 seconds EFPL. We
do not know when the crack formed between these times, for example
cracks may have formed at 1,150 seconds EFPL. (Approximately 2/3 of
the blade sets tested in excess of 1,375 seconds EFPL had cracks. Some
recent experiments have, indeed, shown cracks as early as 1,150
seconds.) It was important to keep the number high, for the Challenger
was to fly an engine very close to the limit by the time the flight
was over.
Finally it is claimed that the criteria are not abandoned, and the
system is safe, by giving up the FAA convention that there should be
no cracks, and considering only a completely fractured blade a
failure. With this definition no engine has yet failed. The idea is
that since there is sufficient time for a crack to grow to a fracture
we can insure that all is safe by inspecting all blades for cracks. If
they are found, replace them, and if none are found we have enough
time for a safe mission. This makes the crack problem not a flight
safety problem, but merely a maintenance problem.
This may in fact be true. But how well do we know that cracks
always grow slowly enough that no fracture can occur in a mission?
Three engines have run for long times with a few cracked blades (about
3,000 seconds EFPL) with no blades broken off.
But a fix for this cracking may have been found. By changing the
blade shape, shot-peening the surface, and covering with insulation to
exclude thermal shock, the blades have not cracked so far.
A very similar story appears in the history of certification of the
HPOTP, but we shall not give the details here.
It is evident, in summary, that the Flight Readiness Reviews and
certification rules show a deterioration for some of the problems of
the Space Shuttle Main Engine that is closely analogous to the
deterioration seen in the rules for the Solid Rocket Booster.
Avionics
By "avionics" is meant the computer system on the Orbiter as well
as its input sensors and output actuators. At first we will restrict
ourselves to the computers proper and not be concerned with the
reliability of the input information from the sensors of temperature,
pressure, etc., nor with whether the computer output is faithfully
followed by the actuators of rocket firings, mechanical controls,
displays to astronauts, etc.
The computer system is very elaborate, having over 250,000 lines of
code. It is responsible, among many other things, for the automatic
control of the entire ascent to orbit, and for the descent until well
into the atmosphere (below Mach 1) once one button is pushed deciding
the landing site desired. It would be possible to make the entire
landing automatically (except that the landing gear lowering signal is
expressly left out of computer control, and must be provided by the
pilot, ostensibly for safety reasons) but such an entirely automatic
landing is probably not as safe as a pilot controlled landing. During
orbital flight it is used in the control of payloads, in displaying
information to the astronauts, and the exchange of information to the
ground. It is evident that the safety of flight requires guaranteed
accuracy of this elaborate system of computer hardware and software.
In brief, the hardware reliability is ensured by having four
essentially independent identical computer systems. Where possible
each sensor also has multiple copies, usually four, and each copy
feeds all four of the computer lines. If the inputs from the sensors
disagree, depending on circumstances, certain averages, or a majority
selection is used as the effective input. The algorithm used by each
of the four computers is exactly the same, so their inputs (since each
sees all copies of the sensors) are the same. Therefore at each step
the results in each computer should be identical. From time to time
they are compared, but because they might operate at slightly
different speeds a system of stopping and waiting at specific times is
instituted before each comparison is made. If one of the computers
disagrees, or is too late in having its answer ready, the three which
do agree are assumed to be correct and the errant computer is taken
completely out of the system. If, now, another computer fails, as
judged by the agreement of the other two, it is taken out of the
system, and the rest of the flight canceled, and descent to the
landing site is instituted, controlled by the two remaining
computers. It is seen that this is a redundant system since the
failure of only one computer does not affect the mission. Finally, as
an extra feature of safety, there is a fifth independent computer,
whose memory is loaded with only the programs of ascent and descent,
and which is capable of controlling the descent if there is a failure
of more than two of the computers of the main line four.
There is not enough room in the memory of the main line computers
for all the programs of ascent, descent, and payload programs in
flight, so the memory is loaded about four time from tapes, by the
astronauts.
Because of the enormous effort required to replace the software for
such an elaborate system, and for checking a new system out, no change
has been made to the hardware since the system began about fifteen
years ago. The actual hardware is obsolete; for example, the memories
are of the old ferrite core type. It is becoming more difficult to
find manufacturers to supply such old-fashioned computers reliably and
of high quality. Modern computers are very much more reliable, can run
much faster, simplifying circuits, and allowing more to be done, and
would not require so much loading of memory, for the memories are much
larger.
The software is checked very carefully in a bottom-up
fashion. First, each new line of code is checked, then sections of
code or modules with special functions are verified. The scope is
increased step by step until the new changes are incorporated into a
complete system and checked. This complete output is considered the
final product, newly released. But completely independently there is
an independent verification group, that takes an adversary attitude to
the software development group, and tests and verifies the software as
if it were a customer of the delivered product. There is additional
verification in using the new programs in simulators, etc. A discovery
of an error during verification testing is considered very serious,
and its origin studied very carefully to avoid such mistakes in the
future. Such unexpected errors have been found only about six times in
all the programming and program changing (for new or altered payloads)
that has been done. The principle that is followed is that all the
verification is not an aspect of program safety, it is merely a test
of that safety, in a non-catastrophic verification. Flight safety is
to be judged solely on how well the programs do in the verification
tests. A failure here generates considerable concern.
To summarize then, the computer software checking system and
attitude is of the highest quality. There appears to be no process of
gradually fooling oneself while degrading standards so characteristic
of the Solid Rocket Booster or Space Shuttle Main Engine safety
systems. To be sure, there have been recent suggestions by management
to curtail such elaborate and expensive tests as being unnecessary at
this late date in Shuttle history. This must be resisted for it does
not appreciate the mutual subtle influences, and sources of error
generated by even small changes of one part of a program on
another. There are perpetual requests for changes as new payloads and
new demands and modifications are suggested by the users. Changes are
expensive because they require extensive testing. The proper way to
save money is to curtail the number of requested changes, not the
quality of testing for each.
One might add that the elaborate system could be very much improved
by more modern hardware and programming techniques. Any outside
competition would have all the advantages of starting over, and
whether that is a good idea for NASA now should be carefully
considered.
Finally, returning to the sensors and actuators of the avionics
system, we find that the attitude to system failure and reliability is
not nearly as good as for the computer system. For example, a
difficulty was found with certain temperature sensors sometimes
failing. Yet 18 months later the same sensors were still being used,
still sometimes failing, until a launch had to be scrubbed because two
of them failed at the same time. Even on a succeeding flight this
unreliable sensor was used again. Again reaction control systems, the
rocket jets used for reorienting and control in flight still are
somewhat unreliable. There is considerable redundancy, but a long
history of failures, none of which has yet been extensive enough to
seriously affect flight. The action of the jets is checked by sensors,
and, if they fail to fire the computers choose another jet to
fire. But they are not designed to fail, and the problem should be
solved.
Conclusions
If a reasonable launch schedule is to be maintained, engineering
often cannot be done fast enough to keep up with the expectations of
originally conservative certification criteria designed to guarantee a
very safe vehicle. In these situations, subtly, and often with
apparently logical arguments, the criteria are altered so that flights
may still be certified in time. They therefore fly in a relatively
unsafe condition, with a chance of failure of the order of a percent
(it is difficult to be more accurate).
Official management, on the other hand, claims to believe the
probability of failure is a thousand times less. One reason for this
may be an attempt to assure the government of NASA perfection and
success in order to ensure the supply of funds. The other may be that
they sincerely believed it to be true, demonstrating an almost
incredible lack of communication between themselves and their working
engineers.
In any event this has had very unfortunate consequences, the most
serious of which is to encourage ordinary citizens to fly in such a
dangerous machine, as if it had attained the safety of an ordinary
airliner. The astronauts, like test pilots, should know their risks,
and we honor them for their courage. Who can doubt that McAuliffe was
equally a person of great courage, who was closer to an awareness of
the true risk than NASA management would have us believe?
Let us make recommendations to ensure that NASA officials deal in a
world of reality in understanding technological weaknesses and
imperfections well enough to be actively trying to eliminate
them. They must live in reality in comparing the costs and utility of
the Shuttle to other methods of entering space. And they must be
realistic in making contracts, in estimating costs, and the difficulty
of the projects. Only realistic flight schedules should be proposed,
schedules that have a reasonable chance of being met. If in this way
the government would not support them, then so be it. NASA owes it to
the citizens from whom it asks support to be frank, honest, and
informative, so that these citizens can make the wisest decisions for
the use of their limited resources.
For a successful technology, reality must take precedence over
public relations, for nature cannot be fooled.
Space Shuttle Challenger Disaster. Date: January 28th, 1986
by R. P. Feynman, 1986.
Richard Feynman With Neil Armstrong at the Rogers Commission Report Press Conference, June 1986.
Introduction
It appears that there are enormous differences of opinion as to the
probability of a failure with loss of vehicle and of human life. The
estimates range from roughly 1 in 100 to 1 in 100,000. The higher
figures come from the working engineers, and the very low figures from
management. What are the causes and consequences of this lack of
agreement? Since 1 part in 100,000 would imply that one could put a
Shuttle up each day for 300 years expecting to lose only one, we could
properly ask "What is the cause of management's fantastic faith in the
machinery?"
We have also found that certification criteria used in Flight
Readiness Reviews often develop a gradually decreasing strictness. The
argument that the same risk was flown before without failure is often
accepted as an argument for the safety of accepting it again. Because
of this, obvious weaknesses are accepted again and again, sometimes
without a sufficiently serious attempt to remedy them, or to delay a
flight because of their continued presence.
There are several sources of information. There are published criteria
for certification, including a history of modifications in the form of
waivers and deviations. In addition, the records of the Flight
Readiness Reviews for each flight document the arguments used to
accept the risks of the flight. Information was obtained from the
direct testimony and the reports of the range safety officer, Louis
J. Ullian, with respect to the history of success of solid fuel
rockets. There was a further study by him (as chairman of the launch
abort safety panel (LASP)) in an attempt to determine the risks
involved in possible accidents leading to radioactive contamination
from attempting to fly a plutonium power supply (RTG) for future
planetary missions. The NASA study of the same question is also
available. For the History of the Space Shuttle Main Engines,
interviews with management and engineers at Marshall, and informal
interviews with engineers at Rocketdyne, were made. An independent
(Cal Tech) mechanical engineer who consulted for NASA about engines
was also interviewed informally. A visit to Johnson was made to gather
information on the reliability of the avionics (computers, sensors,
and effectors). Finally there is a report "A Review of Certification
Practices, Potentially Applicable to Man-rated Reusable Rocket
Engines," prepared at the Jet Propulsion Laboratory by N. Moore, et
al., in February, 1986, for NASA Headquarters, Office of Space
Flight. It deals with the methods used by the FAA and the military to
certify their gas turbine and rocket engines. These authors were also
interviewed informally.
Solid Rockets (SRB)
An estimate of the reliability of solid rockets was made by the range
safety officer, by studying the experience of all previous rocket
flights. Out of a total of nearly 2,900 flights, 121 failed (1 in
25). This includes, however, what may be called, early errors, rockets
flown for the first few times in which design errors are discovered
and fixed. A more reasonable figure for the mature rockets might be 1
in 50. With special care in the selection of parts and in inspection,
a figure of below 1 in 100 might be achieved but 1 in 1,000 is
probably not attainable with today's technology. (Since there are two
rockets on the Shuttle, these rocket failure rates must be doubled to
get Shuttle failure rates from Solid Rocket Booster failure.)
NASA officials argue that the figure is much lower. They point out
that these figures are for unmanned rockets but since the Shuttle is a
manned vehicle "the probability of mission success is necessarily very
close to 1.0." It is not very clear what this phrase means. Does it
mean it is close to 1 or that it ought to be close to 1? They go on to
explain "Historically this extremely high degree of mission success
has given rise to a difference in philosophy between manned space
flight programs and unmanned programs; i.e., numerical probability
usage versus engineering judgment." (These quotations are from "Space
Shuttle Data for Planetary Mission RTG Safety Analysis," Pages 3-1,
3-1, February 15, 1985, NASA, JSC.) It is true that if the probability
of failure was as low as 1 in 100,000 it would take an inordinate
number of tests to determine it ( you would get nothing but a string
of perfect flights from which no precise figure, other than that the
probability is likely less than the number of such flights in the
string so far). But, if the real probability is not so small, flights
would show troubles, near failures, and possible actual failures with
a reasonable number of trials. and standard statistical methods could
give a reasonable estimate. In fact, previous NASA experience had
shown, on occasion, just such difficulties, near accidents, and
accidents, all giving warning that the probability of flight failure
was not so very small. The inconsistency of the argument not to
determine reliability through historical experience, as the range
safety officer did, is that NASA also appeals to history, beginning
"Historically this high degree of mission success..."
Finally, if we are to replace standard numerical probability usage
with engineering judgment, why do we find such an enormous disparity
between the management estimate and the judgment of the engineers? It
would appear that, for whatever purpose, be it for internal or
external consumption, the management of NASA exaggerates the
reliability of its product, to the point of fantasy.
The history of the certification and Flight Readiness Reviews will not
be repeated here. (See other part of Commission reports.) The
phenomenon of accepting for flight, seals that had shown erosion and
blow-by in previous flights, is very clear. The Challenger flight is
an excellent example. There are several references to flights that had
gone before. The acceptance and success of these flights is taken as
evidence of safety. But erosion and blow-by are not what the design
expected. They are warnings that something is wrong. The equipment is
not operating as expected, and therefore there is a danger that it can
operate with even wider deviations in this unexpected and not
thoroughly understood way. The fact that this danger did not lead to a
catastrophe before is no guarantee that it will not the next time,
unless it is completely understood. When playing Russian roulette the
fact that the first shot got off safely is little comfort for the
next. The origin and consequences of the erosion and blow-by were not
understood. They did not occur equally on all flights and all joints;
sometimes more, and sometimes less. Why not sometime, when whatever
conditions determined it were right, still more leading to
catastrophe?
In spite of these variations from case to case, officials behaved as
if they understood it, giving apparently logical arguments to each
other often depending on the "success" of previous flights. For
example. in determining if flight 51-L was safe to fly in the face of
ring erosion in flight 51-C, it was noted that the erosion depth was
only one-third of the radius. It had been noted in an experiment
cutting the ring that cutting it as deep as one radius was necessary
before the ring failed. Instead of being very concerned that
variations of poorly understood conditions might reasonably create a
deeper erosion this time, it was asserted, there was "a safety factor
of three." This is a strange use of the engineer's term ,"safety
factor." If a bridge is built to withstand a certain load without the
beams permanently deforming, cracking, or breaking, it may be designed
for the materials used to actually stand up under three times the
load. This "safety factor" is to allow for uncertain excesses of load,
or unknown extra loads, or weaknesses in the material that might have
unexpected flaws, etc. If now the expected load comes on to the new
bridge and a crack appears in a beam, this is a failure of the
design. There was no safety factor at all; even though the bridge did
not actually collapse because the crack went only one-third of the way
through the beam. The O-rings of the Solid Rocket Boosters were not
designed to erode. Erosion was a clue that something was wrong.
Erosion was not something from which safety can be inferred.
There was no way, without full understanding, that one could have
confidence that conditions the next time might not produce erosion
three times more severe than the time before. Nevertheless, officials
fooled themselves into thinking they had such understanding and
confidence, in spite of the peculiar variations from case to case. A
mathematical model was made to calculate erosion. This was a model
based not on physical understanding but on empirical curve fitting. To
be more detailed, it was supposed a stream of hot gas impinged on the
O-ring material, and the heat was determined at the point of
stagnation (so far, with reasonable physical, thermodynamic laws). But
to determine how much rubber eroded it was assumed this depended only
on this heat by a formula suggested by data on a similar material. A
logarithmic plot suggested a straight line, so it was supposed that
the erosion varied as the .58 power of the heat, the .58 being
determined by a nearest fit. At any rate, adjusting some other
numbers, it was determined that the model agreed with the erosion (to
depth of one-third the radius of the ring). There is nothing much so
wrong with this as believing the answer! Uncertainties appear
everywhere. How strong the gas stream might be was unpredictable, it
depended on holes formed in the putty. Blow-by showed that the ring
might fail even though not, or only partially eroded through. The
empirical formula was known to be uncertain, for it did not go
directly through the very data points by which it was
determined. There were a cloud of points some twice above, and some
twice below the fitted curve, so erosions twice predicted were
reasonable from that cause alone. Similar uncertainties surrounded the
other constants in the formula, etc., etc. When using a mathematical
model careful attention must be given to uncertainties in the model.
Liquid Fuel Engine (SSME)
During the flight of 51-L the three Space Shuttle Main Engines all
worked perfectly, even, at the last moment, beginning to shut down the
engines as the fuel supply began to fail. The question arises,
however, as to whether, had it failed, and we were to investigate it
in as much detail as we did the Solid Rocket Booster, we would find a
similar lack of attention to faults and a deteriorating
reliability. In other words, were the organization weaknesses that
contributed to the accident confined to the Solid Rocket Booster
sector or were they a more general characteristic of NASA? To that end
the Space Shuttle Main Engines and the avionics were both
investigated. No similar study of the Orbiter, or the External Tank
were made.
The engine is a much more complicated structure than the Solid
Rocket Booster, and a great deal more detailed engineering goes into
it. Generally, the engineering seems to be of high quality and
apparently considerable attention is paid to deficiencies and faults
found in operation.
The usual way that such engines are designed (for military or
civilian aircraft) may be called the component system, or bottom-up
design. First it is necessary to thoroughly understand the properties
and limitations of the materials to be used (for turbine blades, for
example), and tests are begun in experimental rigs to determine
those. With this knowledge larger component parts (such as bearings)
are designed and tested individually. As deficiencies and design
errors are noted they are corrected and verified with further
testing. Since one tests only parts at a time these tests and
modifications are not overly expensive. Finally one works up to the
final design of the entire engine, to the necessary
specifications. There is a good chance, by this time that the engine
will generally succeed, or that any failures are easily isolated and
analyzed because the failure modes, limitations of materials, etc.,
are so well understood. There is a very good chance that the
modifications to the engine to get around the final difficulties are
not very hard to make, for most of the serious problems have already
been discovered and dealt with in the earlier, less expensive, stages
of the process.
The Space Shuttle Main Engine was handled in a different manner,
top down, we might say. The engine was designed and put together all
at once with relatively little detailed preliminary study of the
material and components. Then when troubles are found in the
bearings, turbine blades, coolant pipes, etc., it is more expensive
and difficult to discover the causes and make changes. For example,
cracks have been found in the turbine blades of the high pressure
oxygen turbopump. Are they caused by flaws in the material, the effect
of the oxygen atmosphere on the properties of the material, the
thermal stresses of startup or shutdown, the vibration and stresses of
steady running, or mainly at some resonance at certain speeds, etc.?
How long can we run from crack initiation to crack failure, and how
does this depend on power level? Using the completed engine as a test
bed to resolve such questions is extremely expensive. One does not
wish to lose an entire engine in order to find out where and how
failure occurs. Yet, an accurate knowledge of this information is
essential to acquire a confidence in the engine reliability in use.
Without detailed understanding, confidence can not be attained.
A further disadvantage of the top-down method is that, if an
understanding of a fault is obtained, a simple fix, such as a new
shape for the turbine housing, may be impossible to implement without
a redesign of the entire engine.
The Space Shuttle Main Engine is a very remarkable machine. It has
a greater ratio of thrust to weight than any previous engine. It is
built at the edge of, or outside of, previous engineering
experience. Therefore, as expected, many different kinds of flaws and
difficulties have turned up. Because, unfortunately, it was built in
the top-down manner, they are difficult to find and fix. The design
aim of a lifetime of 55 missions equivalent firings (27,000 seconds of
operation, either in a mission of 500 seconds, or on a test stand) has
not been obtained. The engine now requires very frequent maintenance
and replacement of important parts, such as turbopumps, bearings,
sheet metal housings, etc. The high-pressure fuel turbopump had to be
replaced every three or four mission equivalents (although that may
have been fixed, now) and the high pressure oxygen turbopump every
five or six. This is at most ten percent of the original
specification. But our main concern here is the determination of
reliability.
In a total of about 250,000 seconds of operation, the engines have
failed seriously perhaps 16 times. Engineering pays close attention to
these failings and tries to remedy them as quickly as possible. This
it does by test studies on special rigs experimentally designed for
the flaws in question, by careful inspection of the engine for
suggestive clues (like cracks), and by considerable study and
analysis. In this way, in spite of the difficulties of top-down
design, through hard work, many of the problems have apparently been
solved.
A list of some of the problems follows. Those followed by an
asterisk (*) are probably solved:
1.Turbine blade cracks in high pressure fuel turbopumps (HPFTP). (May have been solved.)
2.Turbine blade cracks in high pressure oxygen turbopumps (HPOTP).
3.Augmented Spark Igniter (ASI) line rupture.*
4.Purge check valve failure.*
5.ASI chamber erosion.*
6.HPFTP turbine sheet metal cracking.
7.HPFTP coolant liner failure.*
8.Main combustion chamber outlet elbow failure.*
9.Main combustion chamber inlet elbow weld offset.*
10.HPOTP subsynchronous whirl.*
11.Flight acceleration safety cutoff system (partial failure in a redundant system).*
12.Bearing spalling (partially solved).
13.A vibration at 4,000 Hertz making some engines inoperable, etc.
Many of these solved problems are the early difficulties of a new
design, for 13 of them occurred in the first 125,000 seconds and only
three in the second 125,000 seconds. Naturally, one can never be sure
that all the bugs are out, and, for some, the fix may not have
addressed the true cause. Thus, it is not unreasonable to guess there
may be at least one surprise in the next 250,000 seconds, a
probability of 1/500 per engine per mission. On a mission there are
three engines, but some accidents would possibly be contained, and
only affect one engine. The system can abort with only two
engines. Therefore let us say that the unknown suprises do not, even
of themselves, permit us to guess that the probability of mission
failure do to the Space Shuttle Main Engine is less than 1/500. To
this we must add the chance of failure from known, but as yet
unsolved, problems (those without the asterisk in the list
above). These we discuss below. (Engineers at Rocketdyne, the
manufacturer, estimate the total probability as 1/10,000. Engineers at
marshal estimate it as 1/300, while NASA management, to whom these
engineers report, claims it is 1/100,000. An independent engineer
consulting for NASA thought 1 or 2 per 100 a reasonable estimate.)
The history of the certification principles for these engines is
confusing and difficult to explain. Initially the rule seems to have
been that two sample engines must each have had twice the time
operating without failure as the operating time of the engine to be
certified (rule of 2x). At least that is the FAA practice, and NASA
seems to have adopted it, originally expecting the certified time to
be 10 missions (hence 20 missions for each sample). Obviously the best
engines to use for comparison would be those of greatest total (flight
plus test) operating time -- the so-called "fleet leaders." But what
if a third sample and several others fail in a short time? Surely we
will not be safe because two were unusual in lasting longer. The short
time might be more representative of the real possibilities, and in
the spirit of the safety factor of 2, we should only operate at half
the time of the short-lived samples.
The slow shift toward decreasing safety factor can be seen in many
examples. We take that of the HPFTP turbine blades. First of all the
idea of testing an entire engine was abandoned. Each engine number has
had many important parts (like the turbopumps themselves) replaced at
frequent intervals, so that the rule must be shifted from engines to
components. We accept an HPFTP for a certification time if two samples
have each run successfully for twice that time (and of course, as a
practical matter, no longer insisting that this time be as large as 10
missions). But what is "successfully?" The FAA calls a turbine blade
crack a failure, in order, in practice, to really provide a safety
factor greater than 2. There is some time that an engine can run
between the time a crack originally starts until the time it has grown
large enough to fracture. (The FAA is contemplating new rules that
take this extra safety time into account, but only if it is very
carefully analyzed through known models within a known range of
experience and with materials thoroughly tested. None of these
conditions apply to the Space Shuttle Main Engine.
Cracks were found in many second stage HPFTP turbine blades. In one
case three were found after 1,900 seconds, while in another they were
not found after 4,200 seconds, although usually these longer runs
showed cracks. To follow this story further we shall have to realize
that the stress depends a great deal on the power level. The
Challenger flight was to be at, and previous flights had been at, a
power level called 104% of rated power level during most of the time
the engines were operating. Judging from some material data it is
supposed that at the level 104% of rated power level, the time to
crack is about twice that at 109% or full power level (FPL). Future
flights were to be at this level because of heavier payloads, and many
tests were made at this level. Therefore dividing time at 104% by 2,
we obtain units called equivalent full power level (EFPL). (Obviously,
some uncertainty is introduced by that, but it has not been studied.)
The earliest cracks mentioned above occurred at 1,375 EFPL.
Now the certification rule becomes "limit all second stage blades
to a maximum of 1,375 seconds EFPL." If one objects that the safety
factor of 2 is lost it is pointed out that the one turbine ran for
3,800 seconds EFPL without cracks, and half of this is 1,900 so we are
being more conservative. We have fooled ourselves in three ways. First
we have only one sample, and it is not the fleet leader, for the other
two samples of 3,800 or more seconds had 17 cracked blades between
them. (There are 59 blades in the engine.) Next we have abandoned the
2x rule and substituted equal time. And finally, 1,375 is where we did
see a crack. We can say that no crack had been found below 1,375, but
the last time we looked and saw no cracks was 1,100 seconds EFPL. We
do not know when the crack formed between these times, for example
cracks may have formed at 1,150 seconds EFPL. (Approximately 2/3 of
the blade sets tested in excess of 1,375 seconds EFPL had cracks. Some
recent experiments have, indeed, shown cracks as early as 1,150
seconds.) It was important to keep the number high, for the Challenger
was to fly an engine very close to the limit by the time the flight
was over.
Finally it is claimed that the criteria are not abandoned, and the
system is safe, by giving up the FAA convention that there should be
no cracks, and considering only a completely fractured blade a
failure. With this definition no engine has yet failed. The idea is
that since there is sufficient time for a crack to grow to a fracture
we can insure that all is safe by inspecting all blades for cracks. If
they are found, replace them, and if none are found we have enough
time for a safe mission. This makes the crack problem not a flight
safety problem, but merely a maintenance problem.
This may in fact be true. But how well do we know that cracks
always grow slowly enough that no fracture can occur in a mission?
Three engines have run for long times with a few cracked blades (about
3,000 seconds EFPL) with no blades broken off.
But a fix for this cracking may have been found. By changing the
blade shape, shot-peening the surface, and covering with insulation to
exclude thermal shock, the blades have not cracked so far.
A very similar story appears in the history of certification of the
HPOTP, but we shall not give the details here.
It is evident, in summary, that the Flight Readiness Reviews and
certification rules show a deterioration for some of the problems of
the Space Shuttle Main Engine that is closely analogous to the
deterioration seen in the rules for the Solid Rocket Booster.
Avionics
By "avionics" is meant the computer system on the Orbiter as well
as its input sensors and output actuators. At first we will restrict
ourselves to the computers proper and not be concerned with the
reliability of the input information from the sensors of temperature,
pressure, etc., nor with whether the computer output is faithfully
followed by the actuators of rocket firings, mechanical controls,
displays to astronauts, etc.
The computer system is very elaborate, having over 250,000 lines of
code. It is responsible, among many other things, for the automatic
control of the entire ascent to orbit, and for the descent until well
into the atmosphere (below Mach 1) once one button is pushed deciding
the landing site desired. It would be possible to make the entire
landing automatically (except that the landing gear lowering signal is
expressly left out of computer control, and must be provided by the
pilot, ostensibly for safety reasons) but such an entirely automatic
landing is probably not as safe as a pilot controlled landing. During
orbital flight it is used in the control of payloads, in displaying
information to the astronauts, and the exchange of information to the
ground. It is evident that the safety of flight requires guaranteed
accuracy of this elaborate system of computer hardware and software.
In brief, the hardware reliability is ensured by having four
essentially independent identical computer systems. Where possible
each sensor also has multiple copies, usually four, and each copy
feeds all four of the computer lines. If the inputs from the sensors
disagree, depending on circumstances, certain averages, or a majority
selection is used as the effective input. The algorithm used by each
of the four computers is exactly the same, so their inputs (since each
sees all copies of the sensors) are the same. Therefore at each step
the results in each computer should be identical. From time to time
they are compared, but because they might operate at slightly
different speeds a system of stopping and waiting at specific times is
instituted before each comparison is made. If one of the computers
disagrees, or is too late in having its answer ready, the three which
do agree are assumed to be correct and the errant computer is taken
completely out of the system. If, now, another computer fails, as
judged by the agreement of the other two, it is taken out of the
system, and the rest of the flight canceled, and descent to the
landing site is instituted, controlled by the two remaining
computers. It is seen that this is a redundant system since the
failure of only one computer does not affect the mission. Finally, as
an extra feature of safety, there is a fifth independent computer,
whose memory is loaded with only the programs of ascent and descent,
and which is capable of controlling the descent if there is a failure
of more than two of the computers of the main line four.
There is not enough room in the memory of the main line computers
for all the programs of ascent, descent, and payload programs in
flight, so the memory is loaded about four time from tapes, by the
astronauts.
Because of the enormous effort required to replace the software for
such an elaborate system, and for checking a new system out, no change
has been made to the hardware since the system began about fifteen
years ago. The actual hardware is obsolete; for example, the memories
are of the old ferrite core type. It is becoming more difficult to
find manufacturers to supply such old-fashioned computers reliably and
of high quality. Modern computers are very much more reliable, can run
much faster, simplifying circuits, and allowing more to be done, and
would not require so much loading of memory, for the memories are much
larger.
The software is checked very carefully in a bottom-up
fashion. First, each new line of code is checked, then sections of
code or modules with special functions are verified. The scope is
increased step by step until the new changes are incorporated into a
complete system and checked. This complete output is considered the
final product, newly released. But completely independently there is
an independent verification group, that takes an adversary attitude to
the software development group, and tests and verifies the software as
if it were a customer of the delivered product. There is additional
verification in using the new programs in simulators, etc. A discovery
of an error during verification testing is considered very serious,
and its origin studied very carefully to avoid such mistakes in the
future. Such unexpected errors have been found only about six times in
all the programming and program changing (for new or altered payloads)
that has been done. The principle that is followed is that all the
verification is not an aspect of program safety, it is merely a test
of that safety, in a non-catastrophic verification. Flight safety is
to be judged solely on how well the programs do in the verification
tests. A failure here generates considerable concern.
To summarize then, the computer software checking system and
attitude is of the highest quality. There appears to be no process of
gradually fooling oneself while degrading standards so characteristic
of the Solid Rocket Booster or Space Shuttle Main Engine safety
systems. To be sure, there have been recent suggestions by management
to curtail such elaborate and expensive tests as being unnecessary at
this late date in Shuttle history. This must be resisted for it does
not appreciate the mutual subtle influences, and sources of error
generated by even small changes of one part of a program on
another. There are perpetual requests for changes as new payloads and
new demands and modifications are suggested by the users. Changes are
expensive because they require extensive testing. The proper way to
save money is to curtail the number of requested changes, not the
quality of testing for each.
One might add that the elaborate system could be very much improved
by more modern hardware and programming techniques. Any outside
competition would have all the advantages of starting over, and
whether that is a good idea for NASA now should be carefully
considered.
Finally, returning to the sensors and actuators of the avionics
system, we find that the attitude to system failure and reliability is
not nearly as good as for the computer system. For example, a
difficulty was found with certain temperature sensors sometimes
failing. Yet 18 months later the same sensors were still being used,
still sometimes failing, until a launch had to be scrubbed because two
of them failed at the same time. Even on a succeeding flight this
unreliable sensor was used again. Again reaction control systems, the
rocket jets used for reorienting and control in flight still are
somewhat unreliable. There is considerable redundancy, but a long
history of failures, none of which has yet been extensive enough to
seriously affect flight. The action of the jets is checked by sensors,
and, if they fail to fire the computers choose another jet to
fire. But they are not designed to fail, and the problem should be
solved.
Conclusions
If a reasonable launch schedule is to be maintained, engineering
often cannot be done fast enough to keep up with the expectations of
originally conservative certification criteria designed to guarantee a
very safe vehicle. In these situations, subtly, and often with
apparently logical arguments, the criteria are altered so that flights
may still be certified in time. They therefore fly in a relatively
unsafe condition, with a chance of failure of the order of a percent
(it is difficult to be more accurate).
Official management, on the other hand, claims to believe the
probability of failure is a thousand times less. One reason for this
may be an attempt to assure the government of NASA perfection and
success in order to ensure the supply of funds. The other may be that
they sincerely believed it to be true, demonstrating an almost
incredible lack of communication between themselves and their working
engineers.
In any event this has had very unfortunate consequences, the most
serious of which is to encourage ordinary citizens to fly in such a
dangerous machine, as if it had attained the safety of an ordinary
airliner. The astronauts, like test pilots, should know their risks,
and we honor them for their courage. Who can doubt that McAuliffe was
equally a person of great courage, who was closer to an awareness of
the true risk than NASA management would have us believe?
Let us make recommendations to ensure that NASA officials deal in a
world of reality in understanding technological weaknesses and
imperfections well enough to be actively trying to eliminate
them. They must live in reality in comparing the costs and utility of
the Shuttle to other methods of entering space. And they must be
realistic in making contracts, in estimating costs, and the difficulty
of the projects. Only realistic flight schedules should be proposed,
schedules that have a reasonable chance of being met. If in this way
the government would not support them, then so be it. NASA owes it to
the citizens from whom it asks support to be frank, honest, and
informative, so that these citizens can make the wisest decisions for
the use of their limited resources.
For a successful technology, reality must take precedence over
public relations, for nature cannot be fooled.
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