(a) COMMENTS DUE DATE
The FAA must receive comments on this AD action by February 17, 2015.
(b) AFFECTED ADS
This AD supersedes AD 99-01-05 R1, Amendment 39-17688 (78 FR 79599,
December 31, 2013) "AD 99-01-05 R1". AD 99-26-19 R1, Amendment 39
-17681 (78 FR 76040, December 16, 2013), also relates to the subject
of this AD.
(c) APPLICABILITY
This AD applies to the following airplanes identified in Table 1 and
Table 2 to paragraph (c) of this AD, that are equipped with wing lift
struts, including airplanes commonly known as a "Clipped Wing Cub,"
which modify the airplane primarily by removing approximately 40
inches of the inboard portion of each wing; and are certificated in
any category.
(1) Based on optional engine installations some airplanes may have been
re-identified or registered with another model that is not listed in
the type certificate data sheet (TCDS). For instance, Piper Model J3C
-65 airplanes are type certificated on Type Certificate Data Sheet
(TCDS) A-691 but may also have been re-identified or registered as a
Model J3C-115, J3F-50, J3C-75, J3C-75D, J3C-75S, J3L-75, J3C-85, J3C
-85S, J3C-90, J3F-90, J3F-90S, J3C-100, or J3-L4J airplane.
(2) The airplane model number on the affected airplane or its registry may
or may not contain the dash (-), e.g. J3 and J-3. This AD applies to
both variations.
TABLE 1 TO PARAGRAPH (C) OF THIS AD--AIRPLANES PREVIOUSLY AFFECTED BY AD 99-01-05 R1
|
TYPE CERTIFICATE HOLDER
|
AIRCRAFT MODEL
|
SERIAL NO.
|
| FS 2000 Corp |
L-14 |
All. |
| FS 2001 Corp |
J5A (Army L-4F), J5A-80, J5B (Army L-4G), J5C, AE-1, and HE-1 |
All. |
| FS 2002 Corporation |
PA-14 |
14-1 through 14-523. |
| FS 2003 Corporation |
PA-12 and PA-12S |
12-1 through 12-4036. |
| LAVIA ARGENTINA S.A. (LAVIASA). |
PA-25, PA-25-235, and PA-25-260 |
25-1 through 25-8156024. |
| Piper Aircraft, Inc |
TG-8 (Army TG-8, Navy XLNP-1) |
All. |
| Piper Aircraft, Inc |
E-2 and F-2 |
All. |
| Piper Aircraft, Inc |
J3C-40, J3C-50, J3C-50S, J3C-65 (Army L-4, L-4B, L-4H, L-4J, Navy NE-1 and NE-2), J3C-65S, J3F-50, J3F-50S, J3F-60, J3F-60S, J3F-65 (Army L-4D), J3F-65S, J3L, J3L-S, J3L-65 (Army L-4C), and J3L-65S. |
All. |
| Piper Aircraft, Inc. |
J4, J4A, J4A-S, and J4E (Army L-4E) |
4-401 through 4-1649. |
| Piper Aircraft, Inc |
PA-11 and PA-11S |
11-1 through 11-1678. |
| Piper Aircraft, Inc |
PA-15 |
15-1 through 15-388. |
| Piper Aircraft, Inc |
PA-16 and PA-16S |
16-1 through 16-736. |
| Piper Aircraft, Inc |
PA-17 |
17-1 through 17-215. |
| Piper Aircraft, Inc |
PA-18, PA-18S, PA-18 ‘‘105’’ (Special), PA-18S ‘‘105’’ (Special), PA-18A, PA-18 ‘‘125’’ (Army L-21A), PA-18S ‘‘125’’, PA-18AS ‘‘125’’,PA-18 ‘‘135’’ (Army L-21B), PA-18A ‘‘135’’, PA-18S ‘‘135’’, PA-18AS ‘‘135’’, PA-18 ‘‘150’’, PA-18A ‘‘150’’, PA-18S ‘‘150’’, PA-18AS ‘‘150’’, PA-18A (Restricted), PA-18A ‘‘135’’ (Restricted), and PA-18A ‘‘150’’ (Restricted). |
18-1 through 18-8309025, 18900 through 1809032, and 1809034 through 1809040. |
| Piper Aircraft, Inc |
PA-19 (Army L-18C), and PA-19S |
18-1 through 18-7632 and 19-1, 19-2, and 19-3. |
| Piper Aircraft, Inc |
PA-20, PA-20S, PA-20 ‘‘115’’, PA-20S ‘‘115’’, PA-20 ‘‘135’’, and PA-20S ‘‘135’’ |
20-1 through 20-1121. |
| Piper Aircraft, Inc |
PA-22, PA-22-108, PA-22-135, PA-22S-135, PA-22-150, PA-22S-150, PA-22-160, and PA-22S-160. |
22-1 through 22-9848. |
NOTE 1 TO PARAGRAPH (C) OF THIS AD: There is a serial number overlap
between the Piper PA-18 series airplanes and the Piper Model PA-19 (Army
L-18C) airplanes listed in AD 99-01-05 R1 . Serial numbers 18-1 through
18-7632 listed for the PA-18 series airplanes are also now listed under
Model PA-19 (Army L-18C) and Model PA-19S.
TABLE 2 TO PARAGRAPH (C) OF THIS AD--AIRPLANES NEW TO THIS AD
|
TYPE CERTIFICATE HOLDER
|
AIRCRAFT MODEL
|
SERIAL NO.
|
| Piper Aircraft, Inc. |
J-3 |
1100 through 1200 and 1999 and up that were manufactured before October 15, 1939. |
| Piper Aircraft, Inc. |
J3C-65 (Army L-4A). |
All. |
| Piper Aircraft, Inc. |
J3P |
2325, 2327, 2339, 2340, 2342, 2344, 2345, 2347, 2349, 2351, 2355 and up that were manufactured before January 10, 1942. |
| Piper Aircraft, Inc. |
J4B |
4-400 and up that were manufactured before December 11, 1942. |
| Piper Aircraft, Inc. |
J4F |
4-828 and up. |
(d) SUBJECT
Joint Aircraft System Component (JASC)/Air Transport Association (ATA)
of America Code 57, Wings.
(e) UNSAFE CONDITION
(1) The subject of this AD was originally prompted by reports of corrosion
damage found on the wing lift struts. AD 99-01-05 R1 is being
superseded to include certain Piper Aircraft, Inc. Models J-3, J3C-65
(Army L4A), J3P, J4B, and J4F airplanes that were inadvertently
omitted from the applicability, paragraph (c), of AD 99-01-05 and
subsequently AD 99-01-05 R1. Also, there is a serial number overlap
between Piper Model PA-18s listed in AD 99-01-05 R1 and Piper Model
PA-19 (Army L-18C). Certain serial numbers listed for Model PA-18s are
also listed under Model PA-19 (Army L-18C).
(2) AD 99-01-05 R1 was issued to clarify the FAA's intention that if a
sealed wing lift strut assembly is installed as a replacement part,
the repetitive inspection requirement is terminated only if the seal
is never improperly broken. If the seal is improperly broken, then
that wing lift strut becomes subject to continued repetitive
inspections. We did not intend to promote drilling holes into or
otherwise unsealing a sealed strut. This AD retains all the actions
currently required in AD 99-01-05 R1. There are no new requirements in
this AD except for the addition of certain model airplanes to the
Applicability section of this AD.
(3) We are issuing this AD to detect and correct corrosion and cracking on
the front and rear wing lift struts and forks, which could cause the
wing lift strut to fail. This failure could result in the wing
separating from the airplane.
(f) COMPLIANCE
Unless already done (compliance with AD 99-01-05 R1 and AD 93- 10-06,
Amendment 39-8586 (58 FR 29965, May 25, 1993) "AD 93-010- 06"), do the
following actions within the compliance times specified in paragraphs
(g) through (m) of this AD, including all subparagraphs. Properly
unsealing and resealing a sealed wing lift strut is still considered a
terminating action for the repetitive inspection requirements of this
AD as long as all appropriate regulations and issues are considered,
such as static strength, fatigue, material effects, immediate and
long-term (internal and external) corrosion protection, resealing
methods, etc. Current FAA regulations in 14 CFR 43.13(b) specify that
maintenance performed will result in the part's condition to be at
least equal to its original or properly altered condition. Any
maintenance actions that unseal a sealed wing lift strut should be
coordinated with the Atlanta Aircraft Certification Office (ACO)
through the local airworthiness authority (e.g., Flight Standards
District Office). There are provisions in paragraph (o) of this AD for
approving such actions as an alternative method of compliance (AMOC).
(g) REMOVE WING LIFT STRUTS
(1) For all airplanes previously affected by AD 99-01-05 R1: Within 1
calendar month after February 8, 1999 (the effective date retained
from AD 99-01-05, Amendment 39-10972 (63 FR 72132, December 31, 1998)
"AD 99-01-05"), or within 24 calendar months after the last inspection
done in accordance with AD 93-10-06 (which was superseded by AD 99-01
-05), whichever occurs later, remove the wing lift struts following
Piper Aircraft Corporation Mandatory Service Bulletin (Piper MSB) No.
528D, dated October 19, 1990, or Piper MSB No. 910A, dated October 10,
1989, as applicable. Before further flight after the removal, do the
actions in one of the following paragraphs (h)(1), (h)(2), (i)(1),
(i)(2), or (i)(3) of this AD, including all subparagraphs.
(2) For all airplanes new to this AD (not previously affected by AD 99-01
-05 R1): Within 1 calendar month after the effective date of this AD
or within 24 calendar months after the last inspection done in
accordance with AD 93-10-06 (which was superseded by AD 99-01-05),
whichever occurs later, remove the wing lift struts following Piper
Aircraft Corporation Mandatory Service Bulletin (Piper MSB) No. 528D,
dated October 19, 1990, or Piper MSB No. 910A, dated October 10, 1989,
as applicable. Before further flight after the removal, do the actions
in one of the following paragraphs (h)(1), (h)(2), (i)(1), (i)(2), or
(i)(3) of this AD, including all subparagraphs.
(h) INSPECT WING LIFT STRUTS
For all airplanes listed in this AD: Before further flight after the
removal required in paragraph (g) of this AD, inspect each wing lift
strut following paragraph (h)(1) or (h)(2) of this AD, including all
subparagraphs, or do the wing lift strut replacement following one of
the options in paragraph (i)(1), (i)(2), or (i)(3) of this AD.
(1) Inspect each wing lift strut for corrosion and perceptible dents foll-
owing Piper MSB No. 528D, dated October 19, 1990, or Piper MSB No.
910A, dated October 10, 1989, as applicable.
(i) If no corrosion is visible and no perceptible dents are found on any
wing lift strut during the inspection required in paragraph (h)(1) of
this AD, before further flight, apply corrosion inhibitor to each wing
lift strut following Piper MSB No. 528D, dated October 19, 1990, or
Piper MSB No. 910A, dated October 10, 1989, as applicable.
Repetitively thereafter inspect each wing lift strut at intervals not
to exceed 24 calendar months following the procedures in paragraph
(h)(1) or (h)(2) of this AD, including all subparagraphs.
(ii) If corrosion or perceptible dents are found on any wing lift strut
during the inspection required in paragraph (h)(1) of this AD or
during any repetitive inspection required in paragraph (h)(1)(i) of
this AD, before further flight, replace the affected wing lift strut
with one of the replacement options specified in paragraph (i)(1),
(i)(2), or (i)(3) of this AD. Do the replacement following the
procedures specified in those paragraphs, as applicable.
(2) Inspect each wing lift strut for corrosion following the procedures in
the Appendix to this AD. This inspection must be done by a Level 2 or
Level 3 inspector certified using the guidelines established by the
American Society for Non-destructive Testing or the "Military Standard
for Nondestructive Testing Personnel Qualification and Certification"
(MIL-STD-410E), which can be found on the Internet at
http://aerospacedefense.thomasnet.com/Asset/MIL-STD-410.pdf.
(i) If no corrosion is found on any wing lift strut during the inspection
required in paragraph (h)(2) of this AD and all requirements in the
Appendix to this AD are met, before further flight, apply corrosion
inhibitor to each wing lift strut following Piper MSB No. 528D, dated
October 19, 1990, or Piper MSB No. 910A, dated October 10, 1989, as
applicable. Repetitively thereafter inspect each wing lift strut at
intervals not to exceed 24 calendar months following the procedures in
paragraph (h)(1) or (h)(2) of this AD, including all subparagraphs.
(ii) If corrosion is found on any wing lift strut during the inspection
required in paragraph (h)(2) of this AD or during any repetitive
inspection required in paragraph (h)(2)(i) of this AD, or if any
requirement in the Appendix of this AD is not met, before further
flight after any inspection in which corrosion is found or the
Appendix requirements are not met, replace the affected wing lift
strut with one of the replacement options specified in paragraph
(i)(1), (i)(2), or (i)(3) of this AD. Do the replacement following
the procedures specified in those paragraphs, as applicable.
(i) WING LIFT STRUT REPLACEMENT OPTIONS
Before further flight after the removal required in paragraph (g) of
this AD, replace the wing lift struts following one of the options in
paragraph (i)(1), (i)(2), or (i)(3) of this AD, including all
subparagraphs, or inspect each wing lift strut following paragraph
(h)(1) or (h)(2) of this AD.
(1) Install original equipment manufacturer (OEM) part number wing lift
struts (or FAA-approved equivalent part numbers) that have been
inspected following the procedures in either paragraph (h)(1) or
(h)(2) of this AD, including all subparagraphs, and are found to be
airworthy. Do the installations following Piper MSB No. 528D, dated
October 19, 1990, or Piper MSB No. 910A, dated October 10, 1989, as
applicable. Repetitively thereafter inspect the newly installed wing
lift struts at intervals not to exceed 24 calendar months following
the procedures in either paragraph (h)(1) or (h)(2) of this AD,
including all subparagraphs.
(2) Install new sealed wing lift strut assemblies (or FAA-approved equi-
valent part numbers) (these sealed wing lift strut assemblies also
include the wing lift strut forks) following Piper MSB No. 528D, dated
October 19, 1990, and Piper MSB No. 910A, dated October 10, 1989, as
applicable. Installing one of these new sealed wing lift strut
assemblies terminates the repetitive inspection requirements in
paragraphs (h)(1) and (h)(2) of this AD, and the wing lift strut fork
removal, inspection, and replacement requirement in paragraphs (j) and
(k) of this AD, including all subparagraphs, for that wing lift strut
assembly.
(3) Install F. Atlee Dodge wing lift strut assemblies following F. Atlee
Dodge Aircraft Services, Inc. Installation Instructions No. 3233-I for
Modified Piper Wing Lift Struts Supplemental Type Certificate (STC)
SA4635NM, dated February 1, 1991, which can be found on the Internet
at http://rgl.faa.gov/Regulatory_and_Guidance_Library/rgstc.nsf/0/E726
AAA2831BD20085256CC2000E3DB7?OpenDocument&Highlight=sa4635nm. Repetit-
ively thereafter inspect the newly installed wing lift struts at
intervals not to exceed 60 calendar months following the procedures in
paragraph (h)(1) or (h)(2) of this AD, including all subparagraphs.
(j) REMOVE WING LIFT STRUT FORKS
(1) For all airplanes previously affected by AD 99-01-05 R1, except for
Model PA-25, PA-25-235, and PA-25-260 airplanes: Within the next 100
hours time-in-service (TIS) after February 8, 1999 (the effective date
retained from AD 99-01-05) or within 500 hours TIS after the last
inspection done in accordance with AD 93-10-06 (which was superseded
by AD 99-01-05), whichever occurs later, remove the wing lift strut
forks (unless already replaced in accordance with paragraph (i)(2) of
this AD). Do the removal following Piper MSB No. 528D, dated October
19, 1990, or Piper MSB No. 910A, dated October 10, 1989, as
applicable. Before further flight after the removal, do the actions in
one of the following paragraphs (k) or (l) of this AD, including all
subparagraphs.
(2) For all airplanes new to this AD (not previously affected by AD 99-01-
05 R1): Within the next 100 hours TIS after the effective date of this
AD or within 500 hours TIS after the last inspection done in
accordance with AD 93-10-06 (which was superseded by AD 99-01-05),
whichever occurs later, remove the wing lift strut forks (unless
already replaced in accordance with paragraph (i)(2) of this AD). Do
the removal following Piper MSB No. 528D, dated October 19, 1990, or
Piper MSB No. 910A, dated October 10, 1989, as applicable. Before
further flight after the removal, do the actions in one of the
following paragraphs (k) or (l) of this AD, including all
subparagraphs.
(k) INSPECT AND REPLACE WING LIFT STRUT FORKS
For all airplanes affected by this AD: Before further flight after the
removal required in paragraph (j) of this AD, inspect the wing lift
strut forks following paragraph (k) of this AD, including all
subparagraphs, or do the wing lift strut fork replacement following
one of the options in paragraph (l)(1), (l)(2), (l)(3), or (l)(4) of
this AD, including all subparagraphs. Inspect the wing lift strut
forks for cracks using magnetic particle procedures, such as those
contained in FAA Advisory Circular (AC) 43.13-1B, Chapter 5, which can
be found on the Internet http://rgl.faa.gov/Regulatory_and_Guidance_
Library/rgAdvisoryCircular.nsf/0/99c827db9baac81b86256b4500596c4e/$FIL
E/Chapter%2005.pdf. Repetitively thereafter inspect at intervals not
to exceed 500 hours TIS until the replacement time requirement
specified in paragraph (k)(2) or (k)(3) of this AD is reached provided
no cracks are found.
(1) If cracks are found during any inspection required in paragraph (k) of
this AD or during any repetitive inspection required in paragraph
(k)(2) or (k)(3) of this AD, before further flight, replace the
affected wing lift strut fork with one of the replacement options
specified in paragraph (l)(1), (l)(2), (l)(3), or (l)(4) of this AD,
including all subparagraphs. Do the replacement following the
procedures specified in those paragraphs, as applicable.
(2) If no cracks are found during the initial inspection required in para-
graph (k) of this AD and the airplane is currently equipped with
floats or has been equipped with floats at any time during the
previous 2,000 hours TIS since the wing lift strut forks were
installed, at or before accumulating 1,000 hours TIS on the wing lift
strut forks, replace the wing lift strut forks with one of the
replacement options specified in paragraph (l)(1), (l)(2), (l)(3), or
(l)(4) of this AD, including all subparagraphs. Do the replacement
following the procedures specified in those paragraphs, as applicable.
Repetitively thereafter inspect the newly installed wing lift strut
forks at intervals not to exceed 500 hours TIS following the
procedures specified in paragraph (k) of this AD, including all
subparagraphs.
(3) If no cracks are found during the initial inspection required in para-
graph (k) of this AD and the airplane has never been equipped with
floats during the previous 2,000 hours TIS since the wing lift strut
forks were installed, at or before accumulating 2,000 hours TIS on the
wing lift strut forks, replace the wing lift strut forks with one of
the replacement options specified in paragraph (l)(1), (l)(2), (l)(3),
or (l)(4) of this AD, including all subparagraphs. Do the replacement
following the procedures specified in those paragraphs, as applicable.
Repetitively thereafter inspect the newly installed wing lift strut
forks at intervals not to exceed 500 hours TIS following the
procedures specified in paragraph (k) of this AD, including all
subparagraphs.
(l) WING LIFT STRUT FORK REPLACEMENT OPTIONS
Before further flight after the removal required in paragraph (j) of
this AD, replace the wing lift strut forks following one of the
options in paragraph (l)(1), (l)(2), (l)(3), or (l)(4) of this AD,
including all subparagraphs, or inspect the wing lift strut forks
following paragraph (k) of this AD, including all subparagraphs.
(1) Install new OEM part number wing lift strut forks of the same part
numbers of the existing part (or FAA-approved equivalent part numbers)
that were manufactured with rolled threads. Wing lift strut forks
manufactured with machine (cut) threads are not to be used. Do the
installations following Piper MSB No. 528D, dated October 19, 1990, or
Piper MSB No. 910A, dated October 10, 1989, as applicable.
Repetitively thereafter inspect and replace the newly installed wing
lift strut forks at intervals not to exceed 500 hours TIS following
the procedures specified in paragraph (k) of this AD, including all
subparagraphs.
(2) Install new sealed wing lift strut assemblies (or FAA-approved equi-
valent part numbers) (these sealed wing lift strut assemblies also
include the wing lift strut forks) following Piper MSB No. 528D, dated
October 19, 1990, and Piper MSB No. 910A, dated October 10, 1989, as
applicable. This installation may have already been done through the
option specified in paragraph (i)(2) of this AD. Installing one of
these new sealed wing lift strut assemblies terminates the repetitive
inspection requirements in paragraphs (h)(1) and (h)(2) of this AD,
and the wing lift strut fork removal, inspection, and replacement
requirements in paragraphs (j) and (k) of this AD, including all
subparagraphs, for that wing lift strut assembly.
(3) For the airplanes specified below, install Jensen Aircraft wing lift
strut fork assemblies specified below in the applicable STC following
Jensen Aircraft Installation Instructions for Modified Lift Strut
Fitting. Installing one of these wing lift strut fork assemblies
terminates the repetitive inspection requirement of this AD only for
that wing lift strut fork. Repetitively inspect each wing lift strut
as specified in paragraph (h)(1) or (h)(2) of this AD, including all
subparagraphs.
(i) For Models PA-12 and PA-12S airplanes: STC SA1583NM, which can be
found on the Internet at http://rgl.faa.gov/Regulatory_and_Guidance_
Library/rgstc.nsf/0/2E708575849845B285256CC1008213CA?OpenDocument&High
light=sa1583nm;
(ii) For Model PA-14 airplanes: STC SA1584NM, which can be found on the
Internet at http://rgl.faa.gov/Regulatory_and_Guidance_Library/
rgstc.nsf/0/39872B814471737685256CC1008213D0?OpenDocument&Highlight=s
a1584nm;
(iii) For Models PA-16 and PA-16S airplanes: STC SA1590NM, which can be
found on the Internet at http://rgl.faa.gov/Regulatory_and_Guidance_
Library/rgstc.nsf/0/B28C4162E30D941F85256CC1008213F6?OpenDocument&Hi
ghlight=sa1590nm;
(iv) For Models PA-18, PA-18S, PA-18 "105" (Special), PA-18S "105" (Spec-
ial), PA-18A, PA-18 "125" (Army L-21A), PA-18S "125", PA-18AS "125",
PA-18 "135" (Army L-21B), PA-18A "135", PA-18S "135", PA-18AS "135",
PA-18 "150", PA-18A "150", PA-18S "150", PA-18AS "150", PA-18A
(Restricted), PA-18A "135" (Restricted), and PA-18A "150"
(Restricted ) airplanes: STC SA1585NM, which can be found on
the Internet at http://rgl.faa.gov/Regulatory_and_Guidance_Library/
rgstc.nsf/0/A2BE010FB1CA61A285256CC1008213D6?OpenDocument&Highlight=s
a1585nm;
(v) For Models PA-20, PA-20S, PA-20 "115", PA-20S "115", PA-20 "135", and
PA-20S "135" airplanes: STC SA1586NM, which can be found on the Inter-
net at http://rgl.faa.gov/Regulatory_and_Guidance_Library/rgstc.nsf/0/
873CC69D42C87CF585256CC1008213DC?OpenDocument&Highlight=sa1586nm; and
(vi) For Model PA-22 airplanes: STC SA1587NM, which can be found on the
Internet at http://rgl.faa.gov/Regulatory_and_Guidance_Library/
rgstc.nsf/0/B051D04CCC0BED7E85256CC1008213E0?OpenDocument&Highlight=s
a1587nm.
(4) Install F. Atlee Dodge wing lift strut assemblies following F. Atlee
Dodge Installation Instructions No. 3233-I for Modified Piper Wing
Lift Struts (STC SA4635NM), dated February 1, 1991, which can be found
on the Internet at http://rgl.faa.gov/Regulatory_and_Guidance_Library/
rgstc.nsf/0/E726AAA2831BD20085256CC2000E3DB7?OpenDocument&Highlight=sa
4635nm.
This installation may have already been done in accordance paragraph
(i)(3) of this AD. Installing these wing lift strut assemblies
terminates the repetitive inspection requirements of this AD for the
wing lift strut fork only. Repetitively inspect the wing lift struts
as specified in paragraph (h)(1) or (h)(2) of this AD, including all
subparagraphs.
(m) INSTALL PLACARD
(1) For all airplanes previously affected by AD 99-01-05 R1: Within 1
calendar month after February 8, 1999 (the effective date retained
from AD 99-01-05), or within 24 calendar months after the last
inspection required by AD 93-10-06 (which was superseded by AD 99-01
-05), whichever occurs later, and before further flight after any
replacement of a wing lift strut assembly required by this AD, do one
of the following actions in paragraph (m)(1)(i) or (m)(1)(ii) of this
AD. The "NO STEP" markings required by paragraph (m)(1)(i) or
(m)(1)(ii) of this AD must remain in place for the life of the
airplane.
(i) Install "NO STEP" decal, Piper (P/N) 80944-02, on each wing lift strut
approximately 6 inches from the bottom of the wing lift strut in a way
that the letters can be read when entering and exiting the airplane;
or
(ii) Paint the words "NO STEP" approximately 6 inches from the bottom of
the wing lift strut in a way that the letters can be read when
entering and exiting the airplane. Use a minimum of 1-inch letters
using a color that contrasts with the color of the airplane.
(2) For all airplanes new to this AD (not previously affected by AD 99-01
-05 R1): Within 1 calendar month after the effective date of this AD,
or within 24 calendar months after the last inspection required by AD
93-10-06 (which was superseded by AD 99-01-05), whichever occurs
later, and before further flight after any replacement of a wing lift
strut assembly required by this AD, do one of the following actions in
paragraph (m)(2)(i) or (m)(2)(ii) of this AD. The "NO STEP" markings
required by paragraph (m)(2)(i) or (m)(2)(ii) of this AD must remain
in place for the life of the airplane.
(i) Install "NO STEP" decal, Piper (P/N) 80944-02, on each wing lift strut
approximately 6 inches from the bottom of the wing lift strut in a way
that the letters can be read when entering and exiting the airplane;
or
(ii) Paint the words "NO STEP" approximately 6 inches from the bottom of
the wing lift strut in a way that the letters can be read when
entering and exiting the airplane. Use a minimum of 1-inch letters
using a color that contrasts with the color of the airplane
(n) ALTERNATIVE METHODS OF COMPLIANCE (AMOCS)
(1) The Manager, Atlanta ACO, FAA, has the authority to approve AMOCs for
this AD related to Piper Aircraft, Inc. airplanes; the Manager,
Seattle ACO, FAA has the authority to approve AMOCs for this AD
related to FS 2000 Corp, FS 2001 Corp, FS 2002 Corporation, and FS
2003 Corporation airplanes; and the Manager, Standards Office, FAA,
has the authority to approve AMOCs for this AD related to LAVIA
ARGENTINA S.A. (LAVIASA) airplanes, if requested using the procedures
found in 14 CFR 39.19. In accordance with 14 CFR 39.19, send your
request to your principal inspector or local Flight Standards District
Office, as appropriate. If sending information directly to the manager
of the ACO, send it to the attention of the appropriate person
identified in paragraph (o) of this AD.
(2) Before using any approved AMOC, notify your appropriate principal
inspector, or lacking a principal inspector, the manager of the local
flight standards district office/certificate holding district office.
(3) AMOCs approved for AD 93-10-06, Amendment 39-8586 (58 FR 29965, May
25, 1993), AD 99-01-05, Amendment 39-10972 (63 FR 72132, December 31,
1998), and AD 99-01-05 R1, Amendment 39-17688 (78 FR 79599, December
31, 2013) are approved as AMOCs for this AD.
(o) RELATED INFORMATION
(1) For more information about this AD related to Piper Aircraft, Inc.
airplanes, contact: Gregory "Keith" Noles, Aerospace Engineer, FAA,
Atlanta ACO, 1701 Columbia Avenue, College Park, Georgia 30337; phone:
(404) 474-5551; fax: (404) 474-5606; email: gregory.noles@faa.gov.
(2) For more information about this AD related to FS 2000 Corp, FS 2001
Corp, FS 2002 Corporation, and FS 2003 Corporation airplanes, contact:
Jeff Morfitt, Aerospace Engineer, FAA, Seattle ACO, 1601 Lind Avenue
SW., Renton, Washington 98057; phone: (425) 917-6405; fax: (245) 917
-6590; email: jeff.morfitt@faa.gov.
(3) For more information about this AD related to LAVIA ARGENTINA S.A.
(LAVIASA) airplanes, contact: S.M. Nagarajan, Aerospace Engineer, FAA,
Small Airplane Directorate, 901 Locust, Room 301, Kansas City,
Missouri 64106; telephone: (816) 329-4145; fax: (816) 329-4090; email:
sarjapur.nagarajan@faa.gov.
(4) For service information identified in this AD, contact Piper Aircraft,
Inc., Customer Services, 2926 Piper Drive, Vero Beach, Florida 32960;
telephone: (772) 567-4361; Internet: www.piper.com. Copies of the
instructions to the F. Atlee Dodge STC and information about the
Jensen Aircraft STCs may be obtained from F. Atlee Dodge, Aircraft
Services, LLC., 6672 Wes Way, Anchorage, Alaska 99518-0409, Internet:
www.fadodge.com. You may view this referenced service information at
the FAA, Small Airplane Directorate, 901 Locust, Kansas City, Missouri
64106. For information on the availability of this material at the
FAA, call (816) 329-4148.
APPENDIX TO DOCKET NO. FAA-2014-1083
PROCEDURES AND REQUIREMENTS FOR ULTRASONIC INSPECTION OF PIPER WING LIFT
STRUTS
EQUIPMENT REQUIREMENTS
1. A portable ultrasonic thickness gauge or flaw detector with echo-to-
echo digital thickness readout capable of reading to 0.001-inch and an
A-trace waveform display will be needed to do this inspection.
2. An ultrasonic probe with the following specifications will be needed to
accomplish this inspection: 10 MHz (or higher), 0.283-inch (or smaller)
diameter dual element or delay line transducer designed for thickness
gauging. The transducer and ultrasonic system shall be capable of
accurately measuring the thickness of AISI 4340 steel down to 0.020
-inch. An accuracy of +/- 0.002-inch throughout a 0.020-inch to 0.050
-inch thickness range while calibrating shall be the criteria for
acceptance.
3. Either a precision machined step wedge made of 4340 steel (or similar
steel with equivalent sound velocity) or at least three shim samples of
same material will be needed to accomplish this inspection. One
thickness of the step wedge or shim shall be less than or equal to
0.020-inch, one shall be greater than or equal to 0.050-inch, and at
least one other step or shim shall be between these two values.
4. Glycerin, light oil, or similar non-water based ultrasonic couplants
are recommended in the setup and inspection procedures. Water-based
couplants, containing appropriate corrosion inhibitors, may be
utilized, provided they are removed from both the reference standards
and the test item after the inspection procedure is completed and
adequate corrosion prevention steps are then taken to protect these
items.
NOTE: Couplant is defined as "a substance used between the face of the
transducer and test surface to improve transmission of ultrasonic energy
across the transducer/strut interface."
NOTE: If surface roughness due to paint loss or corrosion is present, the
surface should be sanded or polished smooth before testing to assure a
consistent and smooth surface for making contact with the transducer. Care
shall be taken to remove a minimal amount of structural material. Paint
repairs may be necessary after the inspection to prevent further corrosion
damage from occurring. Removal of surface irregularities will enhance the
accuracy of the inspection technique.
INSTRUMENT SETUP
1. Set up the ultrasonic equipment for thickness measurements as specified
in the instrument's user's manual. Because of the variety of equipment
available to perform ultrasonic thickness measurements, some
modification to this general setup procedure may be necessary. However,
the tolerance requirement of step 13 and the record keeping requirement
of step 14, must be satisfied.
2. If battery power will be employed, check to see that the battery has
been properly charged. The testing will take approximately two hours.
Screen brightness and contrast should be set to match environmental
conditions.
3. Verify that the instrument is set for the type of transducer being
used, i.e. single or dual element, and that the frequency setting is
compatible with the transducer.
4. If a removable delay line is used, remove it and place a drop of
couplant between the transducer face and the delay line to assure good
transmission of ultrasonic energy. Reassemble the delay line transducer
and continue.
5. Program a velocity of 0.231-inch/microsecond into the ultrasonic unit
unless an alternative instrument calibration procedure is used to set
the sound velocity.
6. Obtain a step wedge or steel shims per item 3 of the Equipment
Requirements. Place the probe on the thickest sample using couplant.
Rotate the transducer slightly back and forth to "ring" the transducer
to the sample. Adjust the delay and range settings to arrive at an A
-trace signal display with the first backwall echo from the steel near
the left side of the screen and the second backwall echo near the right
of the screen. Note that when a single element transducer is used, the
initial pulse and the delay line/steel interface will be off of the
screen to the left. Adjust the gain to place the amplitude of the first
backwall signal at approximately 80% screen height on the A-trace.
7. "Ring" the transducer on the thinnest step or shim using couplant.
Select positive half-wave rectified, negative half-wave rectified, or
filtered signal display to obtain the cleanest signal. Adjust the pulse
voltage, pulse width, and damping to obtain the best signal resolution.
These settings can vary from one transducer to another and are also
user dependent.
8. Enable the thickness gate, and adjust the gate so that it starts at the
first backwall echo and ends at the second backwall echo. (Measuring
between the first and second backwall echoes will produce a measurement
of the steel thickness that is not affected by the paint layer on the
strut). If instability of the gate trigger occurs, adjust the gain,
gate level, and/or damping to stabilize the thickness reading.
9. Check the digital display reading and if it does not agree with the
known thickness of the thinnest thickness, follow your instrument's
calibration recommendations to produce the correct thickness reading.
When a single element transducer is used this will usually involve
adjusting the fine delay setting.
10. Place the transducer on the thickest step of shim using couplant.
Adjust the thickness gate width so that the gate is triggered by the
second backwall reflection of the thick section. If the digital
display does not agree with the thickest thickness, follow your
instrument's calibration recommendations to produce the correct
thickness reading. A slight adjustment in the velocity may be
necessary to get both the thinnest and the thickest reading correct.
Document the changed velocity value.
11. Place couplant on an area of the lift strut which is thought to be
free of corrosion and "ring" the transducer to surface. Minor
adjustments to the signal and gate settings may be required to account
for coupling improvements resulting from the paint layer. The
thickness gate level should be set just high enough so as not to be
triggered by irrelevant signal noise. An area on the upper surface of
the lift strut above the inspection area would be a good location to
complete this step and should produce a thickness reading between
0.034-inch and 0.041-inch.
12. Repeat steps 8, 9, 10, and 11 until both thick and thin shim measure-
ments are within tolerance and the lift strut measurement is
reasonable and steady.
13. Verify that the thickness value shown in the digital display is within
+/- 0.002-inch of the correct value for each of the three or more
steps of the setup wedge or shims. Make no further adjustments to the
instrument settings.
14. Record the ultrasonic versus actual thickness of all wedge steps or
steel shims available as a record of setup.
INSPECTION PROCEDURE
1. Clean the lower 18 inches of the wing lift struts using a cleaner that
will remove all dirt and grease. Dirt and grease will adversely affect
the accuracy of the inspection technique. Light sanding or polishing
may also be required to reduce surface roughness as noted in the
Equipment Requirements section.
2. Using a flexible ruler, draw a \1/4\-inch grid on the surface of the
first 11 inches from the lower end of the strut as shown in Piper MSB
No. 528D, dated October 19, 1990, or Piper MSB No. 910A, dated October
10, 1989, as applicable. This can be done using a soft (#2) pencil and
should be done on both faces of the strut. As an alternative to drawing
a complete grid, make two rows of marks spaced every \1/4\-inch across
the width of the strut. One row of marks should be about 11 inches from
the lower end of the strut, and the second row should be several inches
away where the strut starts to narrow. Lay the flexible ruler between
respective tick marks of the two rows and use tape or a rubber band to
keep the ruler in place. See Figure 1.
3. Apply a generous amount of couplant inside each of the square areas or
along the edge of the ruler. Re-application of couplant may be
necessary.
4. Place the transducer inside the first square area of the drawn grid or
at the first \1/4\-inch mark on the ruler and "ring" the transducer to
the strut. When using a dual element transducer, be very careful to
record the thickness value with the axis of the transducer elements
perpendicular to any curvature in the strut. If this is not done, loss
of signal or inaccurate readings can result.
5. Take readings inside each square on the grid or at \1/4\-inch incre-
ments along the ruler and record the results. When taking a thickness
reading, rotate the transducer slightly back and forth and experiment
with the angle of contact to produce the lowest thickness reading
possible. Pay close attention to the A-scan display to assure that the
thickness gate is triggering off of maximized backwall echoes.
NOTE: A reading shall not exceed .041 inch. If a reading exceeds .041
-inch, repeat steps 13 and 14 of the Instrument Setup section before
proceeding further.
6. If the A-trace is unsteady or the thickness reading is clearly wrong,
adjust the signal gain and/or gate setting to obtain reasonable and
steady readings. If any instrument setting is adjusted, repeat steps 13
and 14 of the Instrument Setup section before proceeding further.
7. In areas where obstructions are present, take a data point as close to
the correct area as possible.
NOTE: The strut wall contains a fabrication bead at approximately 40% of
the strut chord. The bead may interfere with accurate measurements in that
specific location.
8. A measurement of 0.024-inch or less shall require replacement of the
strut prior to further flight.
9. If at any time during testing an area is encountered where a valid
thickness measurement cannot be obtained due to a loss of signal
strength or quality, the area shall be considered suspect. These areas
may have a remaining wall thickness of less than 0.020-inch, which is
below the range of this setup, or they may have small areas of
localized corrosion or pitting present. The latter case will result in
a reduction in signal strength due to the sound being scattered from
the rough surface and may result in a signal that includes echoes from
the pits as well as the backwall. The suspect area(s) shall be tested
with a Maule "Fabric Tester" as specified in Piper MSB No. 528D, dated
October 19, 1990, or Piper MSB No. 910A, dated October 10, 1989.
10. Record the lift strut inspection in the aircraft log book.
ILLUSTRATION (Figure 1)
Issued in Kansas City, Missouri, on December 19, 2014. Earl Lawrence,
Manager, Small Airplane Directorate, Aircraft Certification Service.
DATES: We must receive comments on this proposed AD by February 17, 2015.