[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]



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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.

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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

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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