[Federal Register Volume 60, Number 106 (Friday, June 2, 1995)]
[Rules and Regulations]
[Pages 28702-28715]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 95-13468]
=======================================================================
-----------------------------------------------------------------------
DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 39
[Docket No. 92-CE-63-AD; Amendment 39-9251; AD 95-12-01]
Airworthiness Directives; Piper Aircraft Corporation PA-25 Series
Airplanes
AGENCY: Federal Aviation Administration, DOT.
ACTION: Final rule.
-----------------------------------------------------------------------
SUMMARY: This amendment supersedes Airworthiness Directive (AD) 93-21-
12, which currently requires inspecting (one-time visual and dye
penetrant) the wing forward spar fuselage attachment assembly for
cracks or corrosion on certain Piper Aircraft Corporation (Piper) PA-25
series airplanes, and replacing or repairing any cracked or corroded
part. This action requires repetitively inspecting (using ultrasonic
and dye penetrant procedures) the wing forward spar fuselage attachment
assembly for cracks or corrosion, replacing or repairing any cracked or
corroded part, and reporting to the Federal Aviation Administration
(FAA) the results of the inspections. This action is prompted by the
FAA's lack of confidence in detecting internal corrosion in the wing
forward spar fuselage attachment fittings while accomplishing the
inspection methods required by AD 93-21-12. A report of a crack in the
wing forward spar fuselage attachment assembly on an airplane where the
inspection requirements of AD 93-21-12 were accomplished also prompted
this action. The actions specified by this AD are intended to prevent
possible in-flight separation of the wing from the airplane caused by a
cracked or corroded wing forward spar fuselage attachment assembly.
EFFECTIVE DATE: July 7, 1995.
ADDRESSES: Information that applies to this AD may be examined at the
FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558,
601 E. 12th Street, Kansas City, Missouri 64106.
FOR FURTHER INFORMATION CONTACT: Christina Marsh, Aerospace Engineer,
FAA, Atlanta Aircraft Certification Office, Campus Building, 1701
Columbia Avenue, suite 2-160, College Park, Georgia 30337-2748;
telephone (404) 305-7362; facsimile (404) 305-7348.
SUPPLEMENTARY INFORMATION: A proposal to amend part 39 of the Federal
Aviation Regulations (14 CFR part 39) to include an AD that would apply
to Piper PA-25 series airplanes was published in the Federal Register
on January 20, 1995 (60 FR 4119). The action proposed to supersede AD
93-21-12 to require repetitively inspecting (using ultrasonic and dye
penetrant procedures) the wing forward spar fuselage attachment
assembly for cracks or corrosion, and replacing or repairing any
cracked or corroded part. Accomplishment of the proposed actions would
be in accordance with the APPENDIX included at the end of the AD.
Interested persons have been afforded an opportunity to participate
in the making of this amendment. Due consideration has been given to
the comments received.
A number of commenters recommend a longer inspection interval for
the affected airplanes, specifically:
Four commenters recommend that the FAA establish a more
frequent inspection interval for those airplanes operating in
agricultural conditions. Two of the commenters recommend utilizing
the proposed two-year inspection interval for those in agricultural
operation and a longer interval for those in non-agricultural
operation;
One commenter recommends that the repetitive inspection
only apply to those airplanes in agricultural operation;
One commenter recommends a repetitive inspection
interval of 2,000 hours time-in-service (TIS);
Six commenters recommend a 10-year repetitive
inspection interval;
One commenter recommends a 5-year repetitive inspection
interval;
One commenter recommends a 3- to 5-year repetitive
inspection interval for those airplanes in non-agricultural
operation; [[Page 28703]]
One commenter recommends a 5-year repetitive inspection
interval for those in NORMAL category operation; and
One commenter recommends a repetitive inspection
interval of 5 years or 2,000 hours TIS, whichever occurs first.
The FAA analyzed and evaluated all available information relating
to the Piper PA-25 series airplane wing forward spar fuselage
attachment assembly crack and corrosion condition when establishing the
repetitive inspection intervals. Based on this information, no
correlation exists between the type of operation that these airplanes
are utilized and the time it takes for corrosion to develop. The AD
compliance time, including the repetitive inspection interval, is
unchanged as a result of these comments. However, the FAA is adding a
reporting requirement to the final rule as a method of further
analyzing this condition on the PA-25 series airplane fleet. Based on
this data, the FAA may adjust the repetitive inspection interval in the
future.
Three commenters feel that AD action is unjustified because the
Piper PA-25 series airplane design is no different than that of any
other airplane constructed with a steel fuselage frame. While there are
literally thousands of airplanes constructed with steel fuselage
frames, each airplane series or model is unique to its own type design.
AD's are issued to correct an unsafe condition that exists or could
develop on a specific type design aircraft. The FAA continuously
analyzes the data of each specific type design aircraft to determine
whether an unsafe condition exists or could develop for a particular
airplane. Regardless of how many AD's exist on other airplane type
designs utilizing steel fuselage structures, the FAA has received
sufficient data to justify issuing an AD to require repetitive
ultrasonic and dye penetrant inspections of the wing forward spar
fuselage attachment assembly of the Piper PA-25 series airplane type
design. The AD is unchanged as a result of these comments.
Seven commenters feel that there is an increased potential for
causing damage to the airplane during the disassembly and re-assembly
necessary to accomplish the repetitive inspections. The commenters'
main concern is the repeated removal of the close-tolerance attach
bolts every two years. The FAA concurs with the idea that frequent
disassembly and re-assembly of the airplane provides the potential for
damaging the airplane, as is true for removing any component to
facilitate inspection. However, the FAA considers the removal of PA-25
series airplane close-tolerance bolts within the skill requirements of
a mechanic certified in accordance with part 65 of the Federal Aviation
Regulations (14 CFR part 65), and that a mechanic certified in this
manner can assemble and disassemble the airplane in a non-damaging
manner. The AD is unchanged as a result of these comments.
Two commenters state that the probability of wing failure caused by
human error during frequent wing removal is greater than wing failure
caused by a cracked or corroded wing attach fitting. The FAA does not
concur. The FAA has not received any reports, data, or information
related to Piper PA-25 series airplane wing failure caused by
disassembling and reassembling the wing; however, the FAA has received
information and data related to two accidents of Piper PA-25 series
airplanes where the wing failed because of cracked and corroded wing
forward spar fuselage attachment assemblies. The AD is unchanged as a
result of these comments.
Three commenters believe that accomplishing the visual and dye-
penetrant inspections specified in AD 93-21-12 are sufficient to detect
corrosion and cracks in the wing forward spar fuselage attachment
assembly. One commenter states that this assembly may be adequately
inspected without removing the wings. The FAA does not concur. Analysis
of the wing fittings in the two accidents revealed that corrosion
internal to the fitting assembly was a contributing factor to the
failures. The FAA developed the proposed ultrasonic and dye penetrant
inspection procedures while actually examining a Piper PA-25 series
airplane. The development of these procedures confirmed to the FAA that
it is possible to inspect a Piper PA-25 series airplane as required by
AD 93-21-12 and not detect corrosion, and that using ultrasonic
inspection procedures is the only FAA-known way of detecting internal
corrosion in the wing forward spar fuselage attachment assembly on the
affected airplanes. The AD is unchanged as a result of these comments.
Three commenters state that the one-time inspection required by AD
93-21-12 is sufficient. The commenters feel that this AD raised the PA-
25 series airplane operators' awareness of and emphasized to the
applicable mechanics the importance of performing inspections of the
wing forward spar fuselage attachment assembly on a regular basis in
the future. The FAA does not concur. A one-time inspection mandated by
an AD may make airplane operators aware of the importance of future
repetitive inspections; however, AD action mandating ultrasonic and dye
penetrant repetitive inspections is the only method the FAA is aware of
to ensure that the unsafe condition of internal corrosion in the wing
forward spar fuselage attachment assembly on the affected airplanes is
detected and corrected.
One commenter states that the provision for replacing the wing
attach cluster every five years instead of repetitively inspecting
every two years is too short of a repetitive interval. The commenter
feels that, if the existing fittings have been installed for 20 to 30
years, then justification exists for allowing additional time between
repetitive inspections if the cluster is replaced. The FAA partially
concurs. The FAA included this cluster replacement provision to give
owners/operators a grace period if the cluster was recently replaced.
The reason for a five-year threshold is to ensure that repetitive
inspections are initiated on the assembly before corrosion develops or
a crack initiates. The addition of the inspection reporting requirement
will allow the FAA to continuously evaluate this threshold, and, as
appropriate, either extend or shorten the repetitive inspection
interval in the future.
Five commenters believe that repetitive inspections are
unjustified. These commenters state that, because the FAA issued AD 93-
21-12 to require a one-time inspection 20 to 30 years after the PA-25
series airplanes were manufactured, it is unrealistic to believe that
corrosion or cracks could occur in the cluster assembly in the two
years since the initial inspection required by AD 93-21-12. The FAA
does not concur. As stated earlier, the airplanes in the referenced
accidents had corrosion internal to the wing fitting assembly. The FAA
has determined that the inspections currently required by AD 93-21-12
will not adequately detect internal corrosion and, this internal
corrosion could develop to the point of structural failure to the wing
when not inspected ultrasonically on a regular basis. The AD is
unchanged as a result of these comments.
Eleven commenters state that the ultrasonic inspections contained
in the proposal would provide a financial impact upon the operators of
the Piper PA-25 series airplanes. Two of these commenters feel that the
impact could be severe enough to eliminate the Piper PA-25 series
airplane fleet. The FAA concurs that the actions would present a
financial impact upon the Piper PA-25 series airplane operators.
Although the main criteria for issuing an AD is to correct a known
unsafe condition and [[Page 28704]] maintain a level of safety for the
airplane equivalent to that originally certificated, the FAA must
present an estimated cost impact upon the public for each AD. The FAA
analyzes each AD to ensure that the condition specified in the AD is
unsafe and is needed to maintain the original level of safety and that
the estimated cost is a fair representation of reality. The FAA has
determined that the level of safety needed for the Piper PA-25 series
airplanes would no longer be achieved if this AD action was not
mandated, and that the cost presented in the economic paragraph of this
AD is an accurate assessment of the actual cost impact upon the public.
The AD is unchanged as a result of these comments.
One commenter states that the ultrasonic inspection specified in
the proposal is not necessary for the steel fuselage tubing. The FAA
concurs. The requirements of the AD are only to inspect ultrasonically
the wing attach fitting clevis ears for internal corrosion. The AD is
unchanged as a result of this comment.
Two commenters recommend that the FAA include certain corrosion
preventative treatments as an option for extending the time that the
repetitive inspections are required. One of these commenters
specifically recommends packing zinc chromate paste on the wing attach
fitting area or treating the fuselage tubing with linseed oil. The
other commenter recommends treating the clusters with Neutrasol after
the initial inspection to halt any additional corrosion development. At
this time, the FAA does not have enough data to ensure that corrosion
inhibitors will deter or eliminate the development of internal
corrosion of the wing forward spar fuselage attachment assembly. The
FAA will keep these ideas in mind while analyzing the data of the
inspection results obtained through this AD. As in any AD action, the
airplane owners/operators may submit any data or ideas to the FAA as a
request for an alternative method of compliance as specified in
paragraph (k) of the AD. The AD is unchanged as a result of these
comments.
After careful review of all available information related to the
subject presented above, the FAA has determined that air safety and the
public interest require the adoption of the rule as proposed except for
the addition of the reporting requirement and minor editorial
corrections. The FAA has determined that the reporting requirement
addition and the minor editorial corrections will not change the
meaning of the AD over that which was proposed. The addition of the
reporting requirement only adds a paperwork burden upon the public over
that already proposed, and the data obtained from the reports may lead
the FAA to extend the repetitive inspection interval in the future.
The compliance time for this AD is presented in calendar time
instead of hours TIS. The FAA has determined that a calendar time for
compliance is the most desirable method because the unsafe condition
described by this AD is caused by corrosion. Corrosion can occur on
airplanes regardless of whether the airplane is in service or in
storage. Therefore, to ensure that corrosion is detected and corrected
on all airplanes within a reasonable period of time without
inadvertently grounding any airplanes, the FAA is mandating a
compliance schedule based upon calendar time instead of hours TIS.
The FAA estimates that 1,272 airplanes in the U.S. registry will be
affected by this AD, that it will take approximately 30 workhours per
airplane to accomplish the required inspection, and that the average
labor rate is approximately $60 an hour. The FAA has become aware that
the affected airplane owners/operators could incur additional expenses
to have their airplanes ultrasonically inspected. This figure will vary
based on scheduling and travel time; however, for the purposes of this
AD the FAA is using a figure of $500. Based on these figures, the total
cost impact of this AD on U.S. operators is estimated to be $2,925,600.
This figure is based on the assumption that no affected airplane owner/
operator has accomplished the required inspections, and does not
reflect the cost of repetitive inspections. The FAA has no way of
determining how many repetitive inspections a particular owner/operator
may incur. In addition, the figure reflects a $500 expense charge for
the ultrasonic inspection. The FAA anticipates that many of the
affected airplane owners/operators will have ultrasonic expense charges
much less than $500.
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 12612, 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 copy of the final evaluation prepared for this
action is contained in the Rules Docket. A copy of it may be obtained
by contacting the Rules Docket at the location provided under the
caption ADDRESSES.
List of Subjects in 14 CFR Part 39
Air transportation, Aircraft, Aviation safety, Incorporation by
reference, 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. App. 1354(a), 1421 and 1423; 49 U.S.C.
106(g); and 14 CFR 11.89.
Sec. 39.13 [Amended]
2. Section 39.13 is amended by removing AD 93-21-12, Amendment 39-
8763 (58 FR 65104, December 13, 1993), and by adding a new AD to read
as follows:
95-12-01 Piper Aircraft Corporation: Amendment 39-9251; Docket No.
92-CE-63-AD. Supersedes AD 93-21-12, Amendment 39-8763.
Applicability: Models PA-25, PA-25-235, and PA-25-260 airplanes
(all serial numbers), 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 use the authority
provided in paragraph (k) of this AD to request approval from the
FAA. This approval may address either no action, if the current
configuration eliminates the unsafe condition, or different actions
necessary to address the unsafe condition described in this AD. Such
a request should include an assessment of the effect of the changed
configuration on the unsafe condition addressed by this AD. In no
case does the presence of any modification, alteration, or repair
remove any airplane from the applicability of this AD.
[[Page 28705]]
Compliance: Required within the next 12 calendar months after
the effective date of this AD, unless already accomplished, and
thereafter at intervals not to exceed 24 calendar months (except as
noted in paragraph (h) of this AD).
To prevent possible in-flight separation of the wing from the
airplane caused by a cracked or corroded wing forward spar fuselage
attachment assembly, accomplish the following:
(a) Gain access to the left and right wing forward spar fuselage
attach fittings by removing the screws retaining the wing fairing.
Dismantle the wing fillet by removing the screws on the aft edge top
and bottom and removing the wing fairing (see FIGURE 1 of the
Appendix to this AD).
(b) Remove the wing attach bolts and wing. Remove paint from the
wing forward spar fuselage attachment fittings and surrounding
areas; do not sand blast because it may obscure surface indications.
Note 2: Saturation of the bolts with a penetrating oil may
facilitate removal.
(c) Visually inspect the wing forward spar tubular fuselage
attach cluster for damage (cracks, corrosion, rust, or gouges).
Prior to further flight, repair or replace any damaged tubular
member with equivalent material in accordance with FAA Advisory
Circular (AC) No. 43.13-1A, Acceptable Methods, Techniques,
Practices--Aircraft Inspection and Repair.
(d) Inspect (using both dye penetrant and ultrasonic procedures)
the wing forward spar fuselage attach fitting assembly, part numbers
(P/N) 61005-0 (front spar fitting assembly) and 61006-0 (front spar
fitting) for Model PA-25; and P/N 64412-0 (front spar fitting
assembly) and 64003-0 (front spar fitting) for Models PA-25-235 and
PA-25-260, for corrosion and cracks in accordance with the Appendix
to this AD.
(1) If any corrosion is found that meets or exceeds the
parameters presented in the Appendix to this AD or any cracks are
found, prior to further flight, replace the forward spar fuselage
tubular attach cluster with serviceable parts as specified in the
Appendix to this AD.
(2) The inspection procedures in the Appendix of this AD, except
for the dye penetrant inspection procedures, must be accomplished by
a Level 2 inspector certified using the guidelines established by
the American Society for Non-destructive Testing, or MIL-STD-410. A
mechanic with at least an Airframe license may perform the dye
penetrant inspection.
(e) Replacement parts required by this AD shall be of those
referenced and specified in either Figures 3a and 3b, 4a and 4b, or
5a and 5b (as applicable), included as part of the Appendix of this
AD.
(f) Prime and paint all areas where parts were replaced or where
paint is bubbled or gone. Use epoxy paint and primer, and, after
paint has cured, rust inhibit the entire area.
(g) Reinstall all items that were removed.
(h) If a new cluster is installed into the fuselage frame,
repetitive inspections are not required until five years after the
replacement date on the respective fuselage side. This cluster may
be replaced every five years as an alternative to the repetitive
inspections.
(i) Send the results of the inspection required by paragraph (d)
of this AD within 10 calendar days after the inspection to the
Manager, Atlanta Aircraft Certification Office (ACO), Campus
Building, 1701 Columbia Avenue, suite 2-160, College Park, Georgia
30337-2748. Include the airplane model and serial number, the
category of operation the airplane is operated in (normal or
restricted), the location and condition of any cracked or corroded
area, the number of hours TIS of the airplane at the time of
inspection, and the approximate number of hours TIS accrued on the
airplane annually. (Reporting approved by the Office of Management
and Budget under OMB no. 2120-0056.)
(j) 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.
(k) An alternative method of compliance or adjustment of the
initial or repetitive compliance times that provides an equivalent
level of safety may be approved by the Manager, Atlanta Aircraft
Certification Office (ACO), Campus Building, 1701 Columbia Avenue,
suite 2-160, College Park, Georgia 30337-2748. The request shall be
forwarded through an appropriate FAA Maintenance Inspector, who may
add comments and then send it to the Manager, Atlanta ACO.
Note 3: Information concerning the existence of approved
alternative methods of compliance with this AD, if any, may be
obtained from the Atlanta ACO.
(l) The Appendix to this AD may be obtained from the Atlanta ACO
at the address specified in paragraph (k) of this AD. This document
or any other information that relates to this AD may be inspected at
the FAA, Central Region, Office of the Assistant Chief Counsel, Room
1558, 601 E. 12th Street, Kansas City, Missouri.
(m) This amendment (39-9251) supersedes AD 88-11- 05, Amendment
39-5997.
(n) This amendment (39-9251) becomes effective on July 7, 1995.
Appendix to AD 95-12-01--Procedures and Requirements for Wing Forward
Spar Attachment Assembly; Inspection of Piper PA-25 Series Airplanes
Equipment Requirements
1. A portable combination ultrasonic flaw detector with both an
LED thickness readout and an A-trace with thickness gate display.
2. An ultrasonic probe with the following: a 15 MHz 0.25-inch
diameter with a 0.375-inch plastic delay line. An equivalent
permanent delay line transducer that provides adequate sensitivity
and resolution to measure a 0.050-inch steel shim can also be used.
3. Three steel shims within the range of 0.050 to 0.100 inches
are required. To ensure proper calibration, the steel shims should
be smooth and free of dirt. In order to verify the shim thickness,
use a calibrated micrometer to measure the steel shims.
4. Either glycerin, 3-in-1 oil, or equivalent ultrasonic
couplants are used to conduct this test set-up and inspection.
Water-based couplants are not permitted because of the possibility
of initiating long-term corrosion of the wing forward spar fuselage
attachment fittings.
Note: Couplant is defined as ``a substance used between the face
of the transducer and test surface to improve transmission of
ultrasonic energy across this boundary or interface.''
Note: If surface pitting is found on either side of the fitting
ears, lightly sand the surface to obtain a smooth working surface.
Removal of surface irregularities such as pits, rust, scale, and
paint will enhance the accuracy of the inspection technique.
Instrument Calibration
1. Turn the instrument power on and check the battery charge
status. The instrument should have at least 40-percent of available
battery life. The screen brightness and contrast of the display
screen should match the environmental conditions (i.e., outside
sunlight or inside a hangar).
2. Depending on the ultrasonic instrument used, select or verify
the single element transducer setting from the probe selection menu.
If a removable delay line is used, unscrew the plastic delay line
from the transducer. Add couplant to the base of the delay line,
than reattach the delay line.
3. Obtain steel shims with known or measured thickness at or
near 0.050, 0.0075, and 0.100 inches. At least one steel shim shall
be greater than 0.095 inches, one less than or equal to 0.050
inches, and one between these two values. Place the probe on the
thickest steel shim using couplant. Adjust the gain setting to
increase the backwall signal from this steel shim. An A-trace will
appear on the screen and a thickness readout will appear on the
display. The signal on the screen from left to right shows: the
initial pulse, the delay line (the front surface of the steel shim)
and the backwall echo of the steel shim. A second and third multiple
backwall echo may also be seen on the A-trace. Enable the thickness
gate. Adjust the thickness gate to initiate at the delay line to
steel shim interface and terminate at the first backwall echo.
4. Place the probe on the thinnest steel shim using couplant.
Adjust the damping, voltage and pulse width to obtain the maximum
signal response and highest resolution on this steel shim. These
settings can vary from probe to probe and are somewhat dependent on
operator preferences.
5. To stabilize the interface synchronization, adjust the
electronic triggering (blocking gate) to approximately three
quarters of the distance between the initial pulse and the delay
line interface echo. The thickness gate should initiate at the delay
line interface echo and terminate at the first backwall echo.
6. Depending on the instrument and probe, select positive half-
wave rectified signal display or negative half-wave rectified signal
display. This selection should give the best signal display on the
thinnest steel shim. Select the interface synchronization. This
selection automatically starts the thickness gate at the delay time
corresponding to the tip of the plastic delay line.
7. Couple the probe to the thickest steel shim using couplant.
Adjust the range so the [[Page 28706]] A-scan display reads from
0.000 to 0.300 inches. Several multiple backwall echoes will
disappear from the screen.
8. Adjust the thickness gate to trigger on the first return
signal. Of instability of the gate trigger occurs, adjust the gain
and/or damping the stabilize the thickness reading. A thickness
readout should be present on the screen and near the known steel
shim thickness.
9. Adjust the velocity to 0.231 inches/microseconds. The
thickness reading should be the known steel shim thickness. Couple
the transducer to the thinnest steel shim. If the thickness readout
does not agree with the known thickness, adjust the fine delay
setting to produce the known thickness. Re-check the thickest step.
If the readout does not indicate the correct thickness re-adjust the
fine delay setting. After this adjustment is made, record the
thickness values for each of the steel shims on a set-up sheet.
10. Calculate the percent error for each measured steel shim.
The maximum allowable percent error should not exceed 3-percent.
Inspection Procedures
1. Add couplant to the outside inspection surface (Refer to
Figures 3a, 4a and 5a, as applicable). Add the appropriate gain to
obtain the backwall echo from the inspection surface. If the gain
setting is adjusted, re-check the thickness values on the steel
shims. To assure proper coupling to the test sample, twist the probe
clockwise and counter-clockwise (with a 45-degree twist) and
maintain contact with the test surface. During the articulation of
the probe, observe the A-trace on the screen and stop the probe
twist at the point of adequate back surface signal amplitude to
trigger the thickness gate on the first half-cycle. Measure and
record the thickness. Repeat the above process at eight equally-
spaced locations around the surface. The weld bead near the spar
cluster maybe hard to access. Find a suitable location near the weld
and measure the thickness.
2. Add couplant to the inside inspection surface (Refer to
figures 3a, 4a and 5a, as applicable). Add the appropriate gain to
obtain the backwall echo from the inspection surface. To assure
proper coupling to the test sample, twist the probe (clockwise and
counter-clockwise with a 45-degree twist). During the articulation
of the probe, observe the A-trace ion the screen and stop the probe
twist at the point of adequate back surface signal amplitude to
trigger the thickness gate on the first half-cycle. Measure and
record the thickness. Repeat the above process at eight equally-
spaced locations around the surface.
3. If a thickness reading in any one of the eight locations from
paragraph 1 of the Inspection Procedures section (outside section
surface) is .085-inch or less for the PA-25 Model or .055-inch or
less for the PA-25-235 and PA-25-260 Models, or if a thickness
reading in any one of the eight locations from paragraph 2. of the
Inspection Procedures section (inside section surface) is .055-inch
or less for the PA-25 Model or .085-inch or less for the PA-25-235
and PA-25-260 Models, prior to further flight, replace the forward
spar fuselage tubular attach cluster with serviceable parts in
accordance with FAA AC No. 43.13-1A, Acceptable Methods, Techniques,
Practices--Aircraft Inspection and Repair. This procedure requires
the following:
a. Provide for the alignment of the airframe with an appropriate
alignment fixture in accordance with FAA AC No. 43.13-1A, Acceptable
Methods, Techniques, Practices--Aircraft Inspection and Repair.
b. Cut the tubular members as referenced and specified in Figure
2 and either Figures 3a and 3b; Figures 4a and 4b; or Figures 5a and
5b, as applicable.
c. Fabricate a cluster using all applicable part numbers
referenced in Figures 3b, 4b, or 5b, as applicable; and
d. Splice the new cluster into the fuselage frame.
Dye Penetrant Inspection
Inspect the wing forward spar fuselage attach fitting assembly
for cracks using FAA-approved dye penetrant methods. If any cracks
are found, prior to further flight, replace the forward spar
fuselage tubular attach cluster with serviceable parts in accordance
with FAA AC No. 43.13-1A, Acceptable Methods, Techniques,
Practices--Aircraft Inspection and Repair. This procedure requires
the following:
1. Provide for the alignment of the airframe with an appropriate
alignment fixture in accordance with FAA AC No. 43.13-1A, Acceptable
Methods, Techniques, Practices--Aircraft Inspection and Repair.
2. Cut the tubular members as referenced and specified in Figure
2 and either Figures 3a and 3b; Figures 4a and 4b; or Figures 5a and
5b, as applicable.
3. Fabricate a cluster using all applicable part numbers
referenced in Figures 3b, 4b, or 5b, as applicable; and
4. Splice the new cluster into the fuselage frame.
BILLING CODE 4910-13-U
[[Page 28707]]
[GRAPHIC][TIFF OMITTED]TR02JN95.125
[[Page 28708]]
[GRAPHIC][TIFF OMITTED]TR02JN95.126
[[Page 28709]]
[GRAPHIC][TIFF OMITTED]TR02JN95.127
[[Page 28710]]
[GRAPHIC][TIFF OMITTED]TR02JN95.128
[[Page 28711]]
[GRAPHIC][TIFF OMITTED]TR02JN95.129
[[Page 28712]]
[GRAPHIC][TIFF OMITTED]TR02JN95.130
[[Page 28713]]
[GRAPHIC][TIFF OMITTED]TR02JN95.131
[[Page 28714]]
[GRAPHIC][TIFF OMITTED]TR02JN95.132
BILLING CODE 4910-13-C
[[Page 28715]] Issued in Kansas City, Missouri, on May 25, 1995.
Henry A. Armstrong,
Acting Manager, Small Airplane Directorate, Aircraft Certification
Service.
[FR Doc. 95-13468 Filed 6-1-95; 8:45 am]
BILLING CODE 4910-13-U
