[Federal Register Volume 66, Number 208 (Friday, October 26, 2001)]
[Rules and Regulations]
[Pages 54111-54119]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 01-26951]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 39
[Docket No. 95-NM-15-AD; Amendment 39-12485; AD 2001-22-06]
RIN 2120-AA64
Airworthiness Directives; Boeing Model B-17E, F, and G Airplanes
AGENCY: Federal Aviation Administration, DOT.
ACTION: Final rule.
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SUMMARY: This amendment adopts a new airworthiness directive (AD),
applicable to all Boeing Model B-17E, F, and G airplanes, that requires
inspections to detect cracking and corrosion of the wing spar chords,
bolts and bolt holes of the spar chords, and wing terminals; and
correction of any discrepancy found during these inspections. This
amendment is prompted by reports of cracking and corrosion of the wing
spar. The actions specified by this AD are intended to prevent reduced
structural integrity of the wing of the airplane due to the problems
associated with corrosion and cracking of the wing spar.
EFFECTIVE DATE: November 30, 2001.
ADDRESSES: Information concerning this amendment may be obtained from
or examined at the Federal Aviation Administration (FAA), Transport
Airplane Directorate, Rules Docket, 1601 Lind Avenue, SW., Renton,
Washington 98055-4056.
FOR FURTHER INFORMATION CONTACT: James G. Rehrl, Aerospace Engineer,
Airframe Branch, ANM-120S, FAA, Seattle Aircraft Certification Office,
1601 Lind Avenue, SW., Renton, Washington 98055-4056; telephone (425)
227-2783; fax (425) 227-1181.
SUPPLEMENTARY INFORMATION: A proposal to amend part 39 of the Federal
Aviation Regulations (14 CFR part 39) to include an airworthiness
directive (AD) that is applicable to all Boeing Model B-17E, F, and G
airplanes was published in the Federal Register on March 16, 1995 (60
FR 14233). That action proposed to require inspections to detect
cracking and corrosion of the wing spar chords, bolts and bolt holes of
the spar chords, and wing terminals; and correction of any discrepancy
found during these inspections.
Of the approximately 12,600 Boeing Model B-17E, B-17F, and B-17G
bombers produced during World War II, only about a dozen remain in
operation. Since the last B-17 was completed in April 1945, each is now
at least 56 years old. Those remaining are flown primarily in various
forms of airshow displays.
Comments
Interested persons have been afforded an opportunity to participate
in the making of this amendment. Due consideration has been given to
the comments received.
Requests To Withdraw Proposed Rule
Many commenters contend that the proposed AD is unjustified and
that it should be withdrawn accordingly. The commenters present various
reasons for this request.
Several commenters assert that cracking in the spar chord is not a
safety issue because no wing or structural failures, incidents, or
accidents have resulted from the cracking addressed by the proposed AD.
One commenter states that the documented support for the necessity of
the proposed AD (as described in the proposal) is flawed and without
technical or event-based merit. Another states that no proper basis or
need for the issuance of an AD has been established.
Several commenters also refer to B-17s flying with known cracks
without incident, some of which are subject to an unspecified type of
inspection. One commenter notes that cracks were present in some B-17s
during World War II, and limits on the degree of cracking that was
acceptable were described in the Structural Repair Manual. The same
commenter notes that battle damage was corrected with strap or angle
reinforcements. Another commenter reports finding corroded or cracked
spars on several airplanes under major restoration, and on one that ran
off a runway into a ravine, consequently requiring major repairs. The
commenter indicates that, despite the extreme conditions that this
latter airplane encountered, and the implied severity of the spar
cracks, no components failed. One commenter reports inspecting the
cracks on a particular B-17 and noticing surface corrosion in the
cracked area of one B-17. The commenter concludes that since corrosion
takes a period of time (sometimes years) to form, the cracks must have
been there for several years. Another commenter reports that a hairline
crack was observed in the left wing of an airplane in 1979, and that
there has been no change or increase in the size of the crack during
years of subsequent flying. (The commenter did not specify which
structural member contained the crack.) The commenter indicates that a
B-17 engineer indicates that there is no safety problem with hairline
cracks.
The FAA acknowledges that no accidents are known to have occurred
as a result of the conditions addressed by the proposed AD.
Nevertheless, the FAA, as well as the operators, are aware of cracks in
the wing spar chords of certain B-17 airplanes. To date five of the B-
17s either flying or capable of being restored to flight status are
known to have cracks in their wing spar chords. The FAA has determined
that there is no design feature to prevent the crack propagation from
becoming transverse and severing the spar chord. The integrity of this
structure is, therefore, essential for continued safe flight and
landing.
Several commenters point to the service history of the B-17 as
evidence that the proposed actions are not necessary. A few commenters
state that, in proposing this rule, the FAA failed to take into account
the ruggedness of the B-17, and they reference occurrences during World
War II in which some B-17s returned with all four spars broken
[[Page 54112]]
as a result of combat damage. One commenter states that the reason for
the airplanes being able to return safely in spite of the degree of
damage is that 90 percent of the wing strength is in the skin and ribs
of the airplane.
Several commenters justify their requests to withdraw the proposal
based on the fact that the current usage of the airplane is far less
demanding--in terms of weights, altitudes, and environments--than the
conditions encountered during wartime operations. Several commenters
note that none of the subject airplanes fly at gross weight, with most
of them, according to one commenter, flying at 10,000 to 15,000 pounds
under gross. One commenter states that the airplanes subject to the
proposed AD are flown only 50 to 250 hours per year. Additionally, the
commenters assert that the current pilots of these airplanes are more
schooled and proficient than those flying them 50 years ago.
Several commenters also cite the excellent maintenance record on
the B-17s as a reason that the proposal should be withdrawn. They point
out that the subject airplanes are under ``constant surveillance,'' and
are well maintained. The commenters also suggest that the remaining B-
17s are better maintained now than when they were new, with many of
them having been completely restored and many of them being hangared
during the airshow off-season.
The FAA concurs with the commenters' assertions that the subject
airplanes are operating in environments much more favorable than those
encountered during World War II. In addition, the FAA recognizes that,
for the most part, these airplanes are meticulously maintained.
However, the FAA does not concur with the request to withdraw the
proposal based on the conditions in which B-17s operate today, because
such conditions are only partially relevant. Of much greater
significance are the conditions to which any particular airplane has
been exposed over its life-span. While most B-17s may be hangared and
well-maintained now, most, if not all, of the affected airplanes have
been exposed to years of grueling operations such as fire-fighting,
aerial application, etc. Furthermore, even if the airplanes had been
hangared continuously since World War II, moisture could accumulate
from condensation. In fact, most of the subject airplanes have spent
much of their life-span in open storage with no particular protection
from the elements.
One commenter indicates that applicable military technical orders
(the basis to which these aircraft are maintained) allow flights with
known cracks in the spar chord if the cracks meet specified criteria.
The commenter reports that this allowance has been validated by combat
operations, current usage of the airplanes, and the type certificate.
Contrary to the commenters' assertions, continued flight with known
structural defects, such as those addressed by the proposed AD, is
considered a violation of section 91.7 of the Federal Aviation
Regulations (14 CFR 91.7), which requires the pilot in command to
discontinue a flight when an unairworthy structural condition occurs.
The FAA finds that a military technical order written almost 60 years
ago during wartime conditions (when emphasis was placed on short-term
airworthiness risks as opposed to long-term risks such as fatigue and
corrosion) is not an appropriate basis for allowing continued flight
with cracks of this nature. The FAA also is not aware of any specific
FAA approval, either directly or by reference, of a military technical
order that allowed continued flight operations for B-17s with
unrepaired cracked spar chords. In any event, this AD would supersede
such an approval.
One commenter justifies its objection to the proposed rule on the
fact that B-17s are not operated for hire. (The Limited Category type
certification basis prohibits using B-17 airplanes for carriage of
passengers or cargo for compensation or hire.) The FAA infers that the
commenter is implying that a lesser safety standard is therefore
acceptable. The FAA does not concur with the commenter's justification.
The corrective action specified in this AD is needed to ensure the
safety of not only the crew members and any other persons on board, but
also of the many spectators that are in proximity to the affected
airplanes as they participate in airshows.
Several commenters report that removal of the wings requires
significant disassembly and express concern that such removal could
reduce the structural integrity of the spar chord-to-terminal fitting
joints. Two commenters state that it has not been determined that these
cracks reduce the structural integrity of the wing assembly. One
commenter states that replacement of used aircraft hardware with new
hardware will affect the aircraft's ``preset'' and ``harmonics,'' and
may establish a stress concentration, which would reduce the integrity
of the aircraft.
The FAA does not concur. The wings have already been removed and
repaired on at least three B-17 airplanes. The FAA has received no
comments indicating the removal and subsequent reinstallation of the
wings reduced the structural integrity of those airplanes.
Nevertheless, wing removal is not required in all instances, as
discussed below. No change to the final rule in this regard is
necessary.
Clarification of Discussion Section of Proposed Rule
Certain commenters request clarification and correction of language
that appears in the Discussion section of the preamble to the notice of
proposed rulemaking (NPRM). One commenter presents an analysis of the
Discussion section, which includes a number of questions and
suggestions for editorial changes. The commenters specifically request
that the FAA correct certain language related to the description of the
wing spar chord to wing terminal fitting joint. One commenter asks for
clarification regarding the description of the wing spar chord to wing
terminal fitting through bolts being ``seized'' in the joint.
Additionally, the commenters request correction of the discussion of
spar loading that appeared in the NPRM. Additionally, the commenters
pose various questions, such as:
--When was the cracking problem discovered by the FAA?
--On how many airplanes was the cracking discovered?
--How many cracked spars have been found?
--How was the cause of the bolt corrosion and spar chord cracking
attributed to moisture entrapment? Was the moisture accumulation
observed or ``is this a guess?''
The FAA finds that clarification of these issues is necessary. The
commenters note correctly that spar chords mate with the cylindrical,
tapering inner wing attach fitting inserts. Each of the eight joints is
held together by eight close-tolerance bolts. The FAA was informed of
the cracking of the wing spar chord and corrosion of these bolts on
April 26, 1994. One B-17 had been inspected at that time, and
approximately one-third of the 64 bolts in the eight joints were
replaced due to corrosion. At least two bolts had lost almost half of
the cross-sectional area. Some of the eight spar chords were cracked,
and one chord end was broken into pieces. Since receiving that report,
the FAA has learned that cracks have been discovered in the wing spar
chord-to-wing terminal fittings of five of the 12 airplanes either
flying or capable of being restored to flight status.
The FAA notes that cracks have propagated to observed lengths
greater
[[Page 54113]]
than seven inches. As the cracks propagate outboard into the region of
increasing longitudinal tensile and compressive stresses, there is no
design feature to prevent the crack propagation from becoming
transverse and severing the spar chord. Because this area is subject to
high axial loads and this structure is necessary for the continued safe
flight of the airplane, cracking in this area is critical.
Evidence that the bolt corrosion and spar chord cracking were due
to moisture entrapment came from several sources. The first operator to
report this condition found corrosion of the joint bolts and the spar
chords. By design, the spar chord tubes are open at the outboard end,
and the presence of the wing terminal fittings inside the spar chords
traps water at the inboard ends. Cracks known to date run
longitudinally along the spars, which indicates that circumferential
loads are cracking the spars. Pressure from corrosion products between
the spar chord-to-terminal joints would create such circumferential
loads.
Commenters correctly note that the bolts in these fittings are not
seized. Rather, moisture trapped in the inner wing spars has caused
some of the bolts to corrode, which makes removal difficult.
Since the Discussion section of the preamble of an NPRM is not
restated in a final rule, no change to this final rule is necessary in
this regard.
Questions Concerning Applicability of Proposed Rule
One commenter asserts that all B-17 aircraft with large, visible
cracks were built by Douglas, and all had history of damage or severe
use. The commenter states that Vega- and Boeing-built B-17s do not have
a problem with cracking.
The FAA infers from these remarks that the commenter requests that
Vega- and Boeing-built airplanes be excluded from the applicability of
this AD. The FAA does not concur. The FAA notes that, of the
approximately 12,600 Model B-17E, B-17F, and B-17G airplanes produced,
nearly 3,000 were manufactured under license by Douglas, and
approximately 2,750 were manufactured under license by Vega, a
subsidiary of Lockheed.
The dozen or so airplanes still in operation--only about one of
every 1,000 produced--comprise a statistically insignificant sample;
therefore, no conclusions can be drawn statistically from the origin of
the particular airplanes in which cracks have been discovered.
Additionally, the commenter fails to present any evidence, such as
differences in design or production methods, that would suggest
airplanes manufactured by Boeing or Vega are less likely to experience
the unsafe condition addressed by this AD. Further, the FAA is not
aware of any such differences. No change to the applicability of this
final rule is necessary.
Another commenter requests that the applicability of the proposed
AD exclude certain airplanes that have already undergone wing removal,
removal of terminals, replacement of close tolerance bolts, and repair
of spar tubes.
The FAA does not concur that a general exclusion should be made for
those airplanes since the previous actions accomplished on those
airplanes may not provide the necessary level of safety. Operators have
not submitted formal documentation to the FAA describing such previous
actions, and so cannot establish that any actions accomplished
previously on these airplanes definitively meet the criteria of this
AD. In addition, it appears likely that there may be repairs
accomplished previously, such as stop-drilling of cracks found in the
spar chords, that do not adequately address the unsafe condition.
However, paragraph (d) of this final rule provides operators with
the opportunity to present the FAA with data to justify approval of an
inspection or repair accomplished previously as an alternative method
of compliance. This provision enables the FAA to review such
inspections and repairs and determine whether further action is
necessary. Also, NOTE 2 of this AD states specifically that operators
of airplanes on which the terminal fitting-to-spar chord joint was
separated prior to the effective date of this AD, and on which
inspection(s) of and/or repair(s) to the wing terminals-to-spar chords
were accomplished prior to the effective date of this AD, should submit
requests for approval of alternative methods of compliance to the FAA.
Question Concerning Cause of Cracking
Several commenters question whether the cracks have been caused by
corrosion. The commenters state there is no documented proof that
corrosion between the steel wing terminal fitting and the aluminum spar
chord is the cause of the cracking. Several commenters state that the
cracks are due to operational abuses (e.g., heavy landings, operating
above gross weights). Another commenter states that the cracks known to
be present on B-17s have not been attributed to any single cause. That
commenter states that environmental stresses (i.e., temperature changes
between the aluminum spar and the steel trunnion) contributed to the
cracking. One commenter states that moisture accumulation and
consequent corrosion cannot be the cause of the cracking addressed by
this AD because most B-17 owners store their airplanes indoors where
moisture cannot accumulate on the spars. Other commenters suggest that
observed cracking is due to a reported manufacturing procedure in which
the terminal fittings, as well as the spar chord-to-terminal fitting
bolts, were driven into place with hammers.
The FAA clarifies that cracking that has been discovered is not
consistent with the damage that would result from overstresses such as
those suggested by the commenters. However, on the other hand, the
cracking is consistent with the pressure that would result from
products of corrosion in the joints. The FAA finds that the
longitudinal nature of the cracks discovered so far is indicative of
expansion due to corrosion products in the spar chord to terminal
fitting joints. It should be noted that the wing terminal fittings are
steel, while the spar chords are constructed of aluminum. Because steel
and aluminum are dissimilar metals, aluminum will tend to galvanically
corrode if in direct contact with steel, as it is in the B-17 design.
The faying surfaces of these joints have not been the subject of
routine maintenance inspections because of the age of the subject
airplanes.
Requests Concerning Separation of Wing Spar Chord-to-Wing Terminal
Joint
Several commenters indicate that separation of the wing spar chord-
to-wing terminal joint is unnecessary, and that the proposed
requirement to remove all 64 bolts in the eight wing spar chord-to-wing
terminal joints is likewise unnecessary. These commenters offer various
proposals with regard to alternative inspection and repair procedures
and compliance times, which are discussed in the following paragraphs.
Several commenters request that the FAA change the requirements to
remove the most inboard bolt in each wing spar chord joint and to
remove all 64 bolts, as specified in proposed paragraphs (a)(2) and
(b)(2)(i), respectively, so that the three most inboard fasteners in
each joint would not have to be removed. One commenter states that the
most inboard bolt in each of the eight wing spar chord-to-wing terminal
joints should not be removed due to interference with other wing
structure and the fact that the bolt is only \5/8\ inch
[[Page 54114]]
from the end of the spar. Some commenters state that the three most
inboard bolts should not be removed for the reason mentioned previously
(for the most inboard bolt), and because removal of the next two most
inboard bolts would necessitate disassembly of a wing rib to access
those bolts.
The FAA finds that some commenters were apparently misled by the
preamble of the proposed AD as to whether the inspections specified in
paragraph (b) of the AD could be accomplished without actually
separating the wing spar-to-wing terminal joint.
The FAA acknowledges that significant disassembly would be required
to remove the three most inboard bolts on the front and rear spars. The
FAA clarifies that the intent of paragraph (b)(2) of the proposed rule
(designated as paragraph (b)(2)(i) of this final rule) is that the use
of equivalent inspections that do not involve separating the terminal
fitting from the spar chord to detect cracking and corrosion may be
acceptable. The FAA has determined that an acceptable level of safety
can be achieved without removing the three most inboard bolts of a
joint provided: (1) The dye penetrant inspection of the spar-chord
tube-end reveals no cracks; (2) the other five bolts are removed and an
eddy current inspection verifies that the holes are free of cracks; and
(3) a borescope inspection using 10-power magnification reveals that
the first, second, and third most inboard bolts are free of corrosion.
These inspections must be performed on a repetitive basis at 36-month
intervals. Paragraph (b) of this final rule has been reformatted, and
this new alternative procedure is specified in paragraph (b)(2)(ii) of
this AD.
Further, the FAA has made editorial changes to paragraphs
(b)(2)(ii), (b)(2)(ii)(B), and (b)(2)(ii)(C) of the final rule to more
clearly specify which bolts are being referred to in those paragraphs.
In addition, the FAA has determined that the requirement to perform
the high frequency eddy current inspection in accordance with paragraph
(a)(2) of the proposed AD (which included removing the most inboard
bolt during the initial inspection) can be omitted from this AD without
unduly affecting aviation safety, since this inspection is adequately
addressed by paragraph (b) of this AD. Therefore, paragraph (a) of this
AD has been re-structured and re-numbered accordingly.
One commenter that has accomplished extensive repairs on a Model B-
17 airplane in the area that is the subject of this AD states that
separating the terminal fitting from the spar chord is the only method
that will adequately address the unsafe condition (corrosion and
cracking of the wing spar, which could result in reduced structural
integrity of the wing of the airplane). The FAA infers that this
commenter is requesting that the FAA revise paragraph (b)(2) of the
proposed AD to eliminate reference to alternative inspection procedures
that may not include separating the terminal fitting from the spar
chord.
The FAA partially concurs with the commenter's request. The FAA
concurs that inspections that involve separating the terminal fitting
from the spar chord are required for all airplanes with cracks that are
unacceptable for repair. The FAA points out that it also has not
approved any alternative inspection procedures for airplanes that have
no cracks or repairable cracks. The FAA also points out that this AD
does not grant blanket approval for alternative inspection procedures.
All inspections in accordance with this AD are required to be
accomplished in accordance with a method approved by the FAA.
However, the FAA does not concur that inspections must include
separation of the terminal fitting from the spar chord. The FAA finds
that it may be possible, depending on the degree of cracking detected,
for alternative inspection procedures to provide an acceptable level of
safety, even if such procedures do not involve separating the terminal
fitting from the spar chord. Paragraph (b)(2)(i) of this AD specifies
that alternative inspection procedures must meet certain minimum
requirements, which are specified in paragraphs (b)(2)(i)(A),
(b)(2)(i)(B), and (b)(2)(i)(C) of this AD. However, the FAA does not
have the resources to develop these procedures for operators. No change
to the final rule is necessary in this regard.
Requests Concerning Proposed Compliance Time
One commenter requests that the proposed 18-month compliance time
specified in paragraph (b) of the AD be changed to allow the
inspections and any needed repairs to be performed during the winter
months away from the airshow season.
The FAA does not concur that the compliance time should be revised.
An 18-month compliance time, as proposed, does allow compliance during
the winter months; therefore, no change to paragraph (b) is necessary
in that regard.
However, in light of the concern raised by the commenter, the FAA
has determined that the compliance time for the initial inspections
specified in paragraph (a) of this AD can be changed to 180 days
without a significant adverse effect on aviation safety. Paragraph (a)
of this AD has been revised accordingly.
Discussion of Repairs
One commenter suggests that each spar chord should be treated with
corrosion inhibitor after bolt removal, replacement, or remedial
action. The FAA infers that the commenter is requesting that a
requirement for application of corrosion inhibitor be added to
applicable paragraphs in the final rule.
The FAA concurs with the commenter's suggestion that each spar
chord should be treated with corrosion inhibitor. Therefore, paragraph
(a)(2) of the final rule (formerly paragraph (a)(3) in the NPRM) has
been revised to include a requirement for application of a corrosion
inhibitor as suggested. The FAA has determined that such a requirement
will increase the long-term corrosion resistance characteristics of the
affected structure without imposing a significant burden on the
operators of the affected airplanes.
One commenter requests that the FAA require that repairs be
performed in accordance with published repair manuals for the B-17. For
those repairs not covered by a published repair manual, the commenter
believes that repairs should be accomplished with the aid of FAA
Designated Engineering Representatives (DER) or other recognized
experts.
The FAA does not concur. All repairs required by this AD must be
approved by the Manager of the Seattle Aircraft Certification Office,
FAA, regardless of whether those repairs are addressed in a published
B-17 repair manual. Although a World War II-era repair manual may be of
some assistance in that regard, it must be recognized that the value of
such a manual is very limited. The primary concern was short-term
airworthiness; that is, that an airplane was to be repaired
sufficiently to safely complete further combat missions. Long-term
considerations, such as fatigue and corrosion, were secondary.
The FAA also recognizes that there have been considerable advances
in repair and corrosion-prevention practices over the last half-
century. As suggested by the commenter, the FAA encourages review of
any needed repair by an appropriately qualified DER since that would
undoubtedly hasten FAA approval of the repair. (Because the repair
would be related to compliance with an airworthiness directive, a DER
would be authorized only to
[[Page 54115]]
recommend its approval.) However, no change to this final rule is
necessary in this regard.
One commenter requests that the final rule be revised to require
replacement of bolts only ``as needed,'' rather than requiring
replacement of any corroded bolt. The commenter states that it has
accomplished a repair that involved removal of the wings and the
terminal attach fittings. In the course of the repair, approximately
one-third of the wing terminal-to-spar bolts were found to be corroded
to the point where replacement was required. However, the commenter
points out that there was no corrosion of the shear plane of any bolt.
Based on the commenter's statements, the FAA infers that the commenter
is requesting that the final rule be revised to require replacement of
bolts only if corrosion is found at the shear plane area.
The FAA does not concur with the commenter's request. The FAA finds
that it would be inappropriate to allow a bolt found to be corroded to
remain installed on an airplane. The FAA has determined that bolts in
wing spar chord-to-wing terminal joints are critical to the safety of
flight; therefore, those bolts must be free of discrepancies, including
corrosion. In addition to the criticality of the bolts to flight
safety, the bolts must be removed to be inspected fully, and the FAA
has determined that it is more cost effective for operators to replace
corroded bolts with new bolts, rather than to perform frequent
repetitive inspections of corroded, or corrosion-reworked, bolts. No
change to the final rule is necessary in this regard.
In lieu of repairing any cracks found, one commenter requests that
the FAA allow operators to attach 4130 steel straps to the outside of
the wing using existing rivet holes. The FAA does not concur that this
would be an acceptable alternative because steel straps fastened to the
outside of the wings would not provide adequate load paths for the
spar-chord loads. No change to the final rule is necessary in this
regard.
Economic Considerations
Some commenters question the cost impact information presented in
the preamble of the NPRM. These commenters take offense to assumptions
made in that section that ``no operator has yet accomplished any of the
proposed requirements'' and that ``no operator would accomplish those
actions if this AD were not adopted.'' One commenter states that all
owners/operators have already voluntarily undertaken inspections and
repairs as a community. The commenter adds that results of those
inspections revealed that virtually all cracks have been discovered
using detailed visual inspection and non-destructive test (NDT)
inspection methods that did not involve de-mating of the spar/wing
terminal. Other commenters also submit information concerning
previously accomplished inspections and corrective actions.
The FAA finds that clarification of language presented in the cost
impact information of this AD is necessary. The FAA and other federal
agencies are required to propose or adopt a regulation only upon
reasoned determination that the benefits of the intended regulation
justify its costs. The two assumptions mentioned above merely represent
a degree of conservatism taken by the FAA in determining that this AD
will, in fact, be cost effective. They are in no way intended to be
judgmental of what a particular operator would or would not do in the
absence of this AD.
Nevertheless, the FAA has not been provided with specific data
indicating that any of the previously accomplished repairs and
inspections provide the level of safety intended by this AD to
adequately address the identified unsafe condition. It also must be
recognized that, in the absence of an AD, operators of B-17s would not
be required to perform the needed inspections and repairs. If there is
no binding requirement to do so, the statutory responsibility of the
FAA to ensure the safety of the occupants of those airplanes and
persons on the ground watching the airplanes during airshows would not
be fulfilled.
Another commenter states that the statement in the proposal that
indicates the proposed AD ``would not have a significant economic
impact on a substantial number of small entities'' is inconsistent with
the estimated cost of $90,000 per airplane.
The FAA notes that the phrase referenced by the commenter refers to
a statutory requirement imposed by the Regulatory Flexibility Act. That
act is intended to protect small businesses and organizations from
federal rulemaking by requiring agencies to develop and analyze
information concerning the effect of rules on small entities. When the
effects of a rule are likely to be ``significant'' on a ``substantial
number of small entities,'' the agency is expected to take steps that
will reduce the burden. Regarding regulatory flexibility findings in
conjunction with the requirements of ADs, very few ADs will ever reach
the level of having a ``significant economic impact, positive or
negative, on a substantial number of small entities,'' since either
most aircraft operators do not meet the agency's criteria for small
entities, or because the cost of an individual AD usually does not
exceed the agency limit for significant impact (which is $100 million
per year). A statement concerning the impact, or lack of it (as in the
case of this AD), is required to be included in the certification
statement of each AD.
Some commenters state that they cannot afford to separate the wing
spar chord-to-terminal joints to perform the inspection. The commenters
state that the AD, as proposed, would create severe economic hardship,
result in grounding of airplanes, and force the sale of non-flying
airplanes at a financial loss. One commenter acknowledges that the cost
estimates fall within federal guidelines for a rule that is ``not a
significant impact;'' however, the commenter contends that, for the
most part, these airplanes are owned by non-profit organizations that
do not have $90,000 in discretionary funds. Another commenter states
that, under the circumstances, issuance of a precautionary
manufacturer's service bulletin or an FAA Advisory Circular would be
more than adequate.
The FAA recognizes the economic impact of the proposed rule.
However, the FAA notes that an unsafe condition exists in regard to the
integrity of the affected joints, which are essential for safe flight.
The FAA also points out that, as explained previously, paragraph (b)(2)
of the AD provides operators the option of performing an alternative
inspection without separating the joints. The FAA expects that costs
for accomplishment of the alternative inspection will likely be lower
than $90,000 per airplane.
Some commenters believe that the cost of compliance will be much
greater than the estimated $90,000 per airplane. One commenter states
that a consensus of affected owners/operators is that the wing spar/
terminal de-mate would require 2,250 to 2,500 work hours. The commenter
notes that this requirement entails removal and reinstallation of four
engines, and complete de-rigging of the control, electrical wiring,
engine control, and fuel systems. Therefore, the commenter estimates
that costs would be from $125,000 to $150,000 per airplane and a four-
to six-month cessation in aircraft financial support activities.
However, one commenter that has actually performed the alternative
inspections outlined in paragraph (b) of the proposed AD states that
the cost impact was much lower than the estimated amount.
[[Page 54116]]
The FAA finds that no change to the cost impact information, below,
is necessary. The FAA based the cost impact information presented in
this AD on the best data available to date for airplanes on which the
wing spar chord-to-wing terminal fitting separation has been
accomplished. Although the costs may vary somewhat, the actual cost for
a particular airplane is not expected to differ greatly from the
estimated cost of $90,000 per airplane.
Issues Related to Inspection Methods and Procedures
One commenter proposes an alternative to the inspection
requirements of paragraph (a) of the proposed AD. The commenter
suggests that within 90 days, and annually thereafter, a dye penetrant
check be accomplished on the inboard butt of each of the eight spar
tubes. The commenter adds that within 12 months, and tri-annually
thereafter, the interior of all eight spar tubes should be treated with
a moisture and corrosion inhibitor.
The commenter also proposes that within 12 months, and thereafter
at 10-year or 1,000-flight-hour intervals, a detailed inspection
designed to detect cracking in the wing spar tubes or terminal bolt
holes should be accomplished. This inspection would include removal and
inspection of the terminal attach bolts at the fifth and seventh most
inboard locations. The commenter also suggests that operators should
inspect annually and report on the status, migration (or lack thereof),
and condition of any cracks determined to be within acceptable
tolerance criteria.
The FAA does not concur with this request. Since corrosion is
believed to be the cause of the cracking, the FAA finds that the
proposed inspection program at the intervals suggested by the commenter
would not ensure such timely detection of cracking. In addition, an
inspection interval based on flight hours is inappropriate because
damage resulting from corrosion is related to calendar time, not flight
time.
Two commenters indicate that no bolts should be removed during the
inspections required by this proposed AD unless there is obvious damage
to the bolt.
One of these commenters states that no bolts should be removed
``without due cause,'' because the bolts have been in the holes of the
joint for more than 50 years, and molecular transfer will have taken
place between the mating surfaces. The commenter asserts that
replacement of the bolts is likely to cause reduced structural
integrity of the wing terminal-to-spar joints.
Another commenter states that replacement of hardware or parts from
an airplane with new parts or hardware changes the harmonics of the
airplane's vibration frequency and establishes a stress point at the
location of the replacement. This commenter states that the engineers
and master mechanics consulted did not recommend replacement of
hardware unless major damage is detected during a visual inspection,
because the stress created by removal could cause significant damage.
The FAA infers that these commenters are requesting that paragraph
(b) of the proposed rule be revised to eliminate the requirement to
remove the bolts that join the wing terminals to the spar chords.
The FAA does not concur with the request to eliminate the
requirement to remove the bolts. The FAA has determined that performing
inspections of the bolts and bolt holes without removing the bolts does
not ensure that corrosion or cracking would be detected. The FAA finds
that, to ensure the continued safety of the fleet of airplanes, it is
necessary to require at least a one-time removal of five of the bolts
in each joint to inspect the shear planes of the bolts for corrosion
and to inspect the bolt holes for cracks. However, as discussed
previously, the FAA has revised this final rule to allow for the three
inboard fasteners in each joint to remain in place, provided that
certain conditions are met.
One commenter inquires as to what eddy current inspection methods
are approved. The FAA is unaware of any military or industry standards
for eddy current inspections. As stated in paragraph (b)(1) of this AD,
eddy current inspections must be approved by the Manager of the Seattle
Aircraft Certification Office, FAA. The manufacturer, The Boeing
Company, has agreed to allow its specifications to be used for the eddy
current inspections. The FAA has added a new NOTE 9 to this final rule
to indicate that this information is available to operators as needed.
One commenter requests that the FAA develop criteria containing
acceptance/rejection standards of cracks characterized (by the
commenter) as insignificant, monitorable, and unacceptable. The
commenter believes that ``blanket condemnation of any cracking is
unwarranted.''
The FAA does not concur that continued flight with any cracking is
acceptable. As specified in paragraph (c) of this final rule, any
cracking discovered as a result of the required inspections must be
repaired prior to further flight. Such repairs may or may not require
separating the wing spar chord-to-wing terminal joint, depending upon
the severity of the cracking. However, if cracking is found and
repaired without separating the wing spar chord-to-wing terminal joint,
repetitive inspections would be required. The FAA expects that the
operator would propose its inspection program as part of the
documentation needed to secure approval of the proposed repair, in
accordance with paragraph (c) of this AD. Continued crack growth
following repair requires separation of the wing spar chord-to-wing
terminal joint in order to positively address continued cracking
problems. No change to the final rule is required in this regard.
Two commenters question the reference to ``acceptance/rejection
criteria contained in sensitivity level Group IV, MIL-I-25135'' which
is contained in paragraph (a)(1) of the proposed rule. One commenter
notes that the referenced military specification does not contain
acceptance/rejection criteria pertaining to cracks, nor was the
specification intended to do so.
The FAA finds that clarification is necessary. The commenter
correctly notes that the military specification cited in the proposed
rule does not contain acceptance/rejection criteria on cracking. The
FAA clarifies that the intent of paragraph (a)(1) is that the dye
penetrant inspection be performed in accordance with MIL-STD-6866,
using a fluorescent Type 1 penetrant, Method C, Sensitivity Level 3,
inspection. Any cracking that is detected must be repaired in
accordance with a method approved by the Manager of the Seattle
Aircraft Certification Office, FAA, prior to further flight. To
eliminate any confusion in this regard, the wording of paragraph (a)(1)
of this final rule has been revised accordingly.
In addition, the FAA recognizes that a variety of dye penetrant
inspection procedures may be acceptable. Therefore, the FAA has added
Note 4 to the final rule to clarify that operators wanting to use an
alternative procedure may request approval from the Seattle Aircraft
Certification Office per the provision of paragraph (d) of this AD.
Clarification of Visual Inspections
The FAA has revised the final rule to clarify that the type of
visual inspection required by paragraphs (a)(2), (b)(1), and
(b)(2)(i)(C) is a ``detailed visual inspection.'' Further, the
definition of this inspection has been included in a new Note 8 of the
final rule.
[[Page 54117]]
Addition of Other New Notes
The FAA has also revised the final rule to include new Notes 1 and
11:
As a result of communications with the Air Transport Association
(ATA) of America, the FAA has learned that, in general, some operators
may misunderstand the legal effect of ADs on airplanes that are
identified in the applicability provision of the AD, but that have been
altered or repaired in the area addressed by the AD. The FAA points out
that all airplanes identified in the applicability provision of an AD
are legally subject to the AD. If an airplane has been altered or
repaired in the affected area in such a way as to affect compliance
with the AD, the owner or operator is required to obtain FAA approval
for an alternative method of compliance with the AD, in accordance with
the paragraph of each AD that provides for such approvals. Therefore, a
new Note 1 has been added to this final rule to clarify this long-
standing requirement.
In addition, a new Note 11 has been added to the final rule to
strongly encourage owners and operators of the affected airplanes to
coordinate their requests for approvals of alternative methods of
compliance or adjustment of the compliance times pertaining to this AD.
Coordination of a single request (in lieu of a separate request from
each owner/operator) will allow the FAA to more quickly review and
respond.
Conclusion
After careful review of the available data, including the comments
noted above, the FAA has determined that air safety and the public
interest require the adoption of the rule with the changes previously
described. The FAA has determined that these changes will neither
increase the economic burden on any operator nor increase the scope of
the AD.
Cost Impact
There are approximately 12 airplanes of the affected design in the
worldwide fleet. The FAA estimates that 10 airplanes of U.S. registry
will be affected by this AD, that it will take approximately 1,500 work
hours per airplane to accomplish the required actions, and that the
average labor rate is $60 per work hour. Based on these figures, the
cost impact of the AD on U.S. operators is estimated to be $900,000, or
$90,000 per airplane.
The cost impact figure discussed above is based on assumptions that
no operator has yet accomplished any of the requirements of this AD
action, and that no operator would accomplish those actions in the
future if this AD were not adopted. The cost impact figures discussed
in AD rulemaking actions represent only the time necessary to perform
the specific actions actually required by the AD. These figures
typically do not include incidental costs, such as the time required to
gain access and close up, planning time, or time necessitated by other
administrative actions.
Regulatory Impact
The regulations adopted herein will not have substantial direct
effects on the States, on the relationship between the national
government and the States, or on the distribution of power and
responsibilities among the various levels of government. Therefore, in
accordance with Executive Order 13132, it is determined that this final
rule does not have sufficient federalism implications to warrant the
preparation of a Federalism Assessment.
For the reasons discussed above, I certify that this action (1) is
not a ``significant regulatory action'' under Executive Order 12866;
(2) is not a ``significant rule'' under DOT Regulatory Policies and
Procedures (44 FR 11034, February 26, 1979); and (3) will not have a
significant economic impact, positive or negative, on a substantial
number of small entities under the criteria of the Regulatory
Flexibility Act. A final evaluation has been prepared for this action
and it is contained in the Rules Docket. A copy of it may be obtained
from the Rules Docket at the location provided under the caption
ADDRESSES.
List of Subjects in 14 CFR Part 39
Air transportation, Aircraft, Aviation safety, Safety.
Adoption of the Amendment
Accordingly, pursuant to the authority delegated to me by the
Administrator, the Federal Aviation Administration amends part 39 of
the Federal Aviation Regulations (14 CFR part 39) as follows:
PART 39--AIRWORTHINESS DIRECTIVES
1. The authority citation for part 39 continues to read as follows:
Authority: 49 U.S.C. 106(g), 40113, 44701.
Sec. 39.13 [Amended]
2. Section 39.13 is amended by adding the following new
airworthiness directive:
2001-22-06 Boeing: Amendment 39-12485. Docket 95-NM-15-AD.
Applicability: All Model B-17E, F, and G airplanes, certificated
in any category.
Note 1: This AD applies to each airplane identified in the
preceding applicability provision, regardless of whether it has been
modified, altered, or repaired in the area subject to the
requirements of this AD. For airplanes that have been modified,
altered, or repaired so that the performance of the requirements of
this AD is affected, the owner/operator must request approval for an
alternative method of compliance in accordance with paragraph (d) of
this AD. The request should include an assessment of the effect of
the modification, alteration, or repair on the unsafe condition
addressed by this AD; and, if the unsafe condition has not been
eliminated, the request should include specific proposed actions to
address it.
Note 2: For airplanes on which the terminal fitting-to-spar
chord joints were separated prior to the effective date of this AD,
and inspections of and/or repairs to the wing terminals-to-spar
chords were accomplished prior to the effective date of this AD:
Applications for approval of an alternative method of compliance to
the requirements of paragraphs (a) and (b) of this AD must be
submitted to the FAA in accordance with the provisions of paragraph
(d) of this AD.
Compliance: Required as indicated, unless accomplished
previously.
To prevent reduced structural integrity of the wing of the
airplane, accomplish the following:
Inspections and Corrective Actions
(a) Within 180 days after the effective date of this AD,
accomplish the requirements of paragraphs (a)(1) and (a)(2) of this
AD.
(1) Perform a dye penetrant inspection to detect cracking of
each inboard end of the eight aluminum wing spar chords, in
accordance with MIL-STD-6866, using a fluorescent Type 1 penetrant,
Method C, Sensitivity Level 3, inspection. If any cracking is
detected, prior to further flight, repair in accordance with a
method approved by the Manager, Seattle Aircraft Certification
Office (ACO), FAA.
Note 3: The part number (P/N) for the upper wing spar chords is
3-14231-0, and the P/N for the lower wing spar chords is 3-14231-1.
Note 4: Operators desiring to use an alternative dye penetrant
procedure may request approval from the Seattle ACO in accordance
with paragraph (d) of this AD.
Note 5: The following are the P/N's for the terminal fitting-to-
spar chord joint assemblies:
[[Page 54118]]
------------------------------------------------------------------------
Assemblies Assembly part number
------------------------------------------------------------------------
Left Upper Front Spar Joint Assembly... 75-4781-0
Right Upper Front Spar Joint Assembly.. 75-4781-1
Left Lower Front Spar Joint Assembly... 65-4782-512
Right Lower Front Spar Joint Assembly.. 65-4782-513
Left Upper Rear Spar Joint Assembly.... 75-4783-0
Right Upper Rear Spar Joint Assembly... 75-4783-1
Left Lower Rear Spar Joint Assembly.... 75-4784-0
Right Lower Rear Spar Joint Assembly... 75-4784-1
------------------------------------------------------------------------
Note 6: The following are the P/N's for the bolts for the spar
chords:
------------------------------------------------------------------------
Bolts for: Bolt part number
------------------------------------------------------------------------
Upper and Lower Front Spar Chords...... NAS56A36
Upper Rear Spar Chord.................. NAS56A34
Lower Rear Spar Chord.................. NAS56A40-5
------------------------------------------------------------------------
(2) Perform a detailed visual inspection to detect corrosion of
the bolts, as installed, and replace any corroded bolt with a new
bolt having a P/N in the NAS 6606 series in accordance with Army
Technical Order Number 01-20EF-2. Prior to further flight, for all
bolt replacements, accomplish the requirements of paragraphs
(a)(2)(i), (a)(2)(ii), (a)(2)(iii), and (a)(2)(iv) of this AD in
accordance with Army Technical Order Number 01-20EF-2.
Note 7: The following are the P/N's for the replacement bolts
for the spar chords:
------------------------------------------------------------------------
Replacement bolts for: Replacement bolt part number
------------------------------------------------------------------------
Upper and Lower Front Spar............. NAS 6606-51
Upper Rear Spar........................ NAS 6606-47
Lower Rear Spar........................ NAS 6606-56
------------------------------------------------------------------------
Note 8: For the purposes of this AD, a detailed visual
inspection is defined as: ``An intensive visual examination of a
specific structural area, system, installation, or assembly to
detect damage, failure, or irregularity. Available lighting is
normally supplemented with a direct source of good lighting at
intensity deemed appropriate by the inspector. Inspection aids such
as mirror, magnifying lenses, etc., may be used. Surface cleaning
and elaborate access procedures may be required.''
(i) Install a washer having P/N MS 20002C6 under the head of the
bolt, a self-locking nut having P/N NAS 1804-6, and a washer having
P/N MS 200026 under the nut, for each replacement bolt.
(ii) Torque any replacement bolt to 95-105 inch-pounds.
(iii) Oversize replacement bolts by \1/16\ inch, as necessary.
(iv) Apply corrosion inhibiting compound (using BMS 3-23, Type
II or equivalent compound) to the spar chord after bolt removal,
replacement, or other remedial action.
(b) Within 18 months after the effective date of this AD,
accomplish the requirements of either paragraph (b)(1) or (b)(2) of
this AD.
(1) Perform detailed visual and high frequency eddy current
inspections, that include separating all eight wing terminal-to-spar
chord joints, to detect cracking and corrosion of the wing terminals
and spar chords, in accordance with a method approved by the
Manager, Seattle ACO; or
(2) Accomplish either paragraph (b)(2)(i) or (b)(2)(ii) of this
AD.
(i) Perform an equivalent inspection(s) to that required by
paragraph (b)(1) of this AD in accordance with a method approved by
the Manager, Seattle ACO. To be considered acceptable, the
equivalent inspection(s) must include, at a minimum, the criteria
specified in paragraphs (b)(2)(i)(A), (b)(2)(i)(B), and (b)(2)(i)(C)
of this AD.
(A) The inspection must include removal of all 64 bolts that
join the eight wing terminals to the eight spar chords; and
(B) The inspection must adequately detect cracking of the spar
chord, and corrosion between the terminal fitting and the spar
chord; and
(C) The inspection must include a detailed visual inspection to
detect corrosion of the attachment bolts; and a high frequency eddy
current, and borescope inspection at 10-power magnification, of the
bolt holes common to the spar chord-to-wing terminal interface.
(ii) Perform a dye penetrant inspection to detect cracking of
the spar chord tube end; remove the most outboard five bolts in the
joint and perform an eddy current inspection to detect cracking of
the holes; and perform a 10-power magnification borescope inspection
to detect corrosion of the most inboard three bolts. If the criteria
specified in paragraphs (b)(2)(ii)(A), (b)(2)(ii)(B), and
(b)(2)(ii)(C) of this AD are met, removal of the three most inboard
bolts of each terminal-to-spar chord joint is not required. Repeat
the requirements of this paragraph thereafter at intervals not to
exceed 36 months.
(A) Results of the dye penetrant inspection of the spar chord
tube end indicate that there is no cracking; and
(B) Results of the eddy current inspection indicate that the
holes of the five most outboard bolts in the joint are free of
cracks; and
(C) Results of the 10-power magnification borescope inspection
indicate that the most inboard three bolts are free of corrosion.
Note 9: The Boeing Company will make its specifications for eddy
current inspections available to operators as needed.
(c) If any cracking and/or corrosion is detected during any of
the inspections required by paragraphs (a) and (b) of this AD, prior
to further flight, repair in accordance with a method approved by
the Manager, Seattle ACO.
Alternative Methods of Compliance
(d) An alternative method of compliance or adjustment of the
compliance time that provides an acceptable level of safety may be
used if approved by the Manager, Seattle ACO. Operators shall submit
their requests through an appropriate FAA Principal Maintenance
Inspector, who may add comments and then send it to the Manager,
Seattle ACO.
Note 10: Information concerning the existence of approved
alternative methods of compliance with this AD, if any, may be
obtained from the Seattle ACO.
[[Page 54119]]
Note 11: The FAA strongly encourages owners and operators of the
affected airplanes to coordinate their requests for approvals of
alternative methods of compliance or adjustment of the compliance
times pertaining to this AD. Coordination of a single request (in
lieu of a separate request from each owner/operator) will allow the
FAA to more quickly review and respond.
Special Flight Permits
(e) Special flight permits may be issued in accordance with
sections 21.197 and 21.199 of the Federal Aviation Regulations (14
CFR 21.197 and 21.199) to operate the airplane to a location where
the requirements of this AD can be accomplished.
Effective Date
(f) This amendment becomes effective on November 30, 2001.
Issued in Renton, Washington, on October 19, 2001.
Ali Bahrami,
Acting Manager, Transport Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 01-26951 Filed 10-25-01; 8:45 am]
BILLING CODE 4910-13-U