[Federal Register Volume 62, Number 145 (Tuesday, July 29, 1997)]
[Rules and Regulations]
[Pages 40702-40706]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 97-19040]
[[Page 40701]]
_______________________________________________________________________
Part VI
Department of Transportation
_______________________________________________________________________
Federal Aviation Administration
_______________________________________________________________________
14 CFR Part 25
Revised Structural Loads Requirements for Transport Category Airplanes;
Final Rule
Federal Register / Vol. 62, No. 145 / Tuesday, July 29, 1997 / Rules
and Regulations
[[Page 40702]]
DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 25
[Docket No. 28312; Amdt. No. 25-91]
RIN 2120-AF70
Revised Structural Loads Requirements for Transport Category
Airplanes
AGENCY: Federal Aviation Administration (FAA), DOT.
ACTION: Final rule.
-----------------------------------------------------------------------
SUMMARY: This amendment revises the structural loads design
requirements of the Federal Aviation Regulations (FAR) for transport
category airplanes by incorporating changes developed in cooperation
with the Joint Aviation Authorities (JAA) of Europe and the Aviation
Rulemaking Advisory Committee (ARAC). This action makes some of the
requirements more rational and eliminates differences between current
U.S. and European requirements that impose unnecessary costs on
airplane manufacturers. These changes are intended to achieve common
airworthiness standards and language between the requirements of the
U.S. regulations and the Joint Aviation Requirements (JAR) of Europe
while maintaining at least the level of safety provided by the current
regulations and industry practices.
EFFECTIVE DATE: August 28, 1997.
FOR FURTHER INFORMATION CONTACT: James Haynes, Airframe and Propulsion
Branch, ANM-112, Transport Airplane Directorate, Aircraft Certification
Service, FAA, 1601 Lind Avenue, SW., Renton, WA 98055-4056; telephone
(206) 227-2131.
SUPPLEMENTARY INFORMATION:
Background
The manufacturing, marketing and certification of transport
airplanes is increasingly an international endeavor. In order for U.S.
manufacturers to export transport airplanes to other countries the
airplane must be designed to comply, not only with the U.S.
airworthiness requirements for transport airplanes (14 CFR part 25),
but also with the transport airworthiness requirements of the countries
to which the airplane is to be exported, unless the importing country
accepts the aircraft without findings of compliance with specified
regulations.
The European countries have developed a common airworthiness code
for transport category airplanes that is administered by the JAA. This
code is the result of a European effort to harmonize the various
airworthiness codes of the European countries and is called the Joint
Aviation Requirements (JAR)-25. It was developed in a format similar to
14 CFR part 25. Many other countries have airworthiness codes that are
aligned closely to part 25 or to JAR-25, or they use these codes
directly for their own certification purposes.
Although JAR-25 is very similar to part 25, there are differences
in methodologies and criteria that often result in the need to address
the same design objective with more than one kind of analysis or test
in order to satisfy both part 25 and JAR airworthiness codes. These
differences result in additional costs to the transport airplane
manufacturers and additional costs to the U.S. and foreign authorities
that must continue to monitor compliance with different airworthiness
codes.
In 1988, the FAA, in cooperation with the JAA and other
organizations representing the U.S. and European aerospace industries,
began a process to harmonize the airworthiness requirements of the
United States and the European authorities. The objective was to
achieve common requirements for the certification of transport category
airplanes without a substantive change in the level of safety provided
by the regulations and industry practices. Other airworthiness
authorities such as Transport Canada have also participated in this
process.
In 1992, the harmonization effort was undertaken by the Aviation
Rulemaking Advisory Committee (ARAC). A working group of industry and
government structural loads specialists of Europe, the United States,
and Canada was chartered by notice in the Federal Register (58 FR
13819, March 15, 1993) to harmonize the design loads sections of
Subpart C of part 25. The bulk of the harmonization tasks for Subpart C
were completed by the working group and recommendations were submitted
to FAA by letter dated February 2, 1995. The FAA concurred with the
recommendations and proposed them in Notice of Proposed Rulemaking
(NPRM) No. 95-14; which was published in the Federal Register on August
29, 1995 (60 FR 44998).
In establishing a design requirement for the nose gear, its
attaching structure and the forward fuselage structure, Sec. 25.499(e)
continues to require consideration of positioning the nose gear in any
steerable position. The term ``any'' is continued from the current
regulation. The term, and the requirements of the section, are
understood in the engineering and regulated communities to require
demonstration that the nose gear and associated structures will sustain
the applicable loads throughout the full range of nose gear positions.
Discussion of Comments
Comments were received from transport airplane manufacturers,
industry associations and foreign airworthiness authorities. All of the
commenters express support for the proposals in Notice No. 95-14
although a few make some recommendations for changes. One comment
believes the changes proposed for Sec. 25.415 could be a burden to some
applicants with airplanes that are derived from models that were
certified to earlier amendment levels of the FAR and JAR. To provide
relief for these derivative airplanes, the commenter proposes a change
to paragraph (b) of Sec. 25.415 which would allow the use of
``realistic'' aerodynamic hinge moment coefficients for control
surfaces in lieu of the prescribed coefficients of paragraph (b). The
FAA does not agree that there is likely to be a burden for derivative
airplanes since the proposed rule applies to new designs. In addition,
the design gust speed does not create an increased requirement over
existing design requirements. Part 24 and JAR-25 were identical in
using 88 feet per second (about 52 knots) in defining hinge moment for
ground gust conditions. However, JAR Sec. 25.519 prescribes a 65 knot
wind speed for ground gusts during jacking and tie-down, and
specifically requires application of those gusts to control surfaces.
As a result, aircraft designs already have to meet the 65 knot rather
than the 52 knot requirement. The ARAC recommends, with FAA and JAA
concurrence, that ground gusts on control surfaces be addressed in just
one section, Sec. 25.415, so Notice No. 95-14 proposes to revise this
section to achieve the same effect as the Sec. 25.519 of JAR-25 by
incorporating the 65-knot wind speed into Sec. 25.415. The net effect
is that there is no change in the ground gust speed requirement for
control surfaces over that already required by JAR-25.
Furthermore, the use of rational aerodynamic hinge moment
coefficients would necessitate a rational ground gust speed as well,
and the 65 knot design gust speed is not necessarily a rational design
speed for ground gusts. Jet blasts in airport operations and normal
storm conditions often exceed 65 knots but service history has shown
that the 65 knot design speed when combined with the conservative
prescribed hinge moments of paragraph (b) provides a satisfactory
design.
[[Page 40703]]
One commenter recommends that the formulation of the requirement
for hinge moments in Sec. 25.415 be changed to show the 65 knot wind
speed explicitly rather than embedding this value into the multiplying
constant. The FAA agrees that this has merit since the connection
between the 65 knot wind speed of Secs. 25.415 and 25.519 could
otherwise be missed in any future rulemaking actions. The rule is
adopted with a change to show the 65 knot wind speed explicitly in the
formula for control surface hinge moments.
One commenter points out that the proposed revision to paragraph
(a) of Sec. 25.481 references paragraphs 25.479(c)(1) and (2) for
vertical and drag load conditions and that these latter paragraphs, as
proposed, no longer specify those conditions. Notice 95-14 proposes to
express the substance of Sec. 25.479(c)(1) and (2) in more general
terms in Sec. 25.473(c). The commenter is correct. The rule is adopted
with a change to delete the incorrect references.
Regulatory Evaluation Summaries
Regulatory Evaluation, Regulatory Flexibility Determination, and Trade
Impact Assessment
Changes to Federal regulations must undergo several economic
analyses. First, Executive Order 12866 directs that each Federal agency
shall propose or adopt a regulation only upon a reasoned determination
that the benefits of the intended regulation justify its costs. Second,
the Regulatory Flexibility Act of 1980 requires agencies to analyze the
economic effect of regulatory changes on small entities. Third, the
Office of Management and Budget directs agencies to assess the effects
of regulatory changes on international trade. In conducting these
analyses, the FAA has determined that this rule:
(1) Will generate benefits that justify its costs and is not a
``significant regulatory action'' as defined in the Executive Order;
(2) is not significant as defined in DOT's Regulatory Policies and
Procedures; (3) will not have a significant impact on a substantial
number of small entities; and (4) will not constitute a barrier to
international trade. These analyses, available in the docket, are
summarized below.
Regulatory Evaluation Summary
Depending on airplane design, the rule could result in additional
compliance costs for some manufacturers. If manufacturers choose to
design to and justify a VD-VC magin of 0.05 Mach,
there will be an increase in analysis costs of approximately $145,000
per certification. The requirement in Sec. 25.473 to consider
structural flexibility in the analysis of landing loads and the
increase in the factor on the maximum static reaction on the nose gear
vertical force in Sec. 25.499 could add compliance costs, but the FAA
estimates that these will be negligible.
The rule will also result in cost savings. Revisions in the
conditions in which unchecked pitch maneuvers are investigated could
reduce certification costs by as much as $10,000 per certification. The
FAA estimates that the change in the speed margin between VB
and VC from a fixed margin to a margin variable with
altitude could result in substantial, though unquantified, cost savings
to some manufacturers. Manufacturers that design small transport
category airplanes with direct mechanical rudder control systems could
realize a savings as a result of the modification in the rudder control
force limit in Sec. 25.351. No comments were received on the costs or
cost savings resulting from these changes.
The primary benefit of the rule will be the cost savings associated
with harmonization of the FAR with the JAR. In order to sell airplanes
in a global marketplace, manufacturers usually certify their products
under the FAR and the JAR. The cost savings from reducing the resources
necessary to demonstrate compliance with non-harmonized design load
requirements will outweigh any incremental costs of the rule, resulting
in a net cost savings. These savings will be realized by U.S.
manufacturers that market airplanes in JAA countries as well as by
manufacturers in JAA countries that market airplanes in the U.S.
The change to Sec. 25.335(b)(2) in the minimum speed margin for
atmospheric conditions from 0.05 Mach to 0.07 Mach could produce safety
benefits. The increase in the margin between VD/
MD and VC/MC is more conservative and
will standardize training across international lines. Crews could
cross-train and cross-fly and this standardization will enhance safety
as well as result in more efficient training.
Regulatory Flexibility Determination
The Regulatory Flexibility Act of 1980 (RFA) was enacted by
Congress to ensure that small entities are not unnecessarily and
disproportionally burdened by Federal regulations. The RFA requires a
Regulatory Flexibility Analysis if a proposed or final rule would have
a significant economic impact, either detrimental or beneficial, on a
substantial number of small entities. FAA Order 2100.14A, Regulatory
Flexibility Criteria and Guidance, establishes threshold cost values
and small entity standards for complying with RFA review requirements
in FAA rulemaking actions. The Order defines ``small entities'' in
terms of size threshold, ``significant economic impact'' in terms of
annualized cost thresholds, and ``substantial number'' as a number
which is not less than eleven and which is more than one-third of the
small entities subject to the proposed or final rule.
Order 2100.14A specifies a size threshold for classification as a
small manufacturer as 75 or fewer employees. Since none of the
manufacturers affected by this rule has 75 or fewer employees and any
costs of the rule will be negligible, the rule will not have a
significant economic impact on a substantial number of small
manufacturers.
International Trade Impact Assessment
The rule will not constitute a barrier to international trade,
including the export of U.S. airplanes to foreign markets and the
import of foreign airplanes into the U.S. Because the rule will
harmonize with the JAR, it would, in fact, lessen restraints on trade.
Federalism Implications
The regulations amended herein do not have a 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. Thus, in
accordance with Executive Order 12612, it is determined that this rule
does not have sufficient federalism implications to warrant the
preparation of a Federalism Assessment.
International Compatibility
In keeping with U.S. obligations under the Convention on
International Civil Aviation, it is FAA policy to comply with
International Civil Aviation Organization (ICAO) standards and
recommended practices to the maximum extent practicable. The FAA has
determined that this rule does not conflict with any international
agreement of the United States.
Paperwork Reduction Act
In accordance with the Paperwork Reduction Act of 1980 (Pub. L. 96-
511), there are no requirements for information collection associated
with this rule.
[[Page 40704]]
Conclusion
Because these changes to the structural loads requirements do not
result in any substantial economic costs, the FAA has determined that
this rule will not be significant under Executive Order 12866. Because
there has not been significant public interest in this issue, the FAA
has determined that this action is not significant under DOT Regulatory
Policies and Procedures (44 FR 11034; February 25, 1979). In addition,
since there are no small entities affected by this rulemaking, the FAA
certifies that the rule 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, since none will be
affected. A copy of the regulatory evaluation prepared for this project
may be examined in the Rules Docket or obtained from the person
identified under the caption FOR FURTHER INFORMATION CONTACT.
List of Subjects in 14 CFR Part 25
Air transportation, Aircraft, Aviation safety, Safety.
The Amendments
Accordingly, the Federal Aviation Administration (FAA) amends 14
CFR part 25 of the Federal Aviation Regulations as follows:
PART 25--AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES
1. The authority citation for part 25 continues to read as follows:
Authority: 49 U.S.C. 106(g), 40113, 44701-44702, 44704.
2. Section 25.331 is amended by revising the introductory text of
paragraph (c) and paragraph (c)(1) to read as follows:
Sec. 25.331 Symmetric maneuvering conditions.
* * * * *
(c) Pitch maneuver conditions. The conditions specified in
paragraphs (c)(1) and (2) of this section must be investigated. The
movement of the pitch control surfaces may be adjusted to take into
account limitations imposed by the maximum pilot effort specified by
Sec. 25.397(b), control system stops and any indirect effect imposed by
limitations in the output side of the control system (for example,
stalling torque or maximum rate obtainable by a power control system.)
(1) Maximum pitch control displacement at VA. The
airplane is assumed to be flying in steady level flight (point
A1, Sec. 25.333(b)) and the cockpit pitch control is
suddenly moved to obtain extreme nose up pitching acceleration. In
defining the tail load, the response of the airplane must be taken into
account. Airplane loads that occur subsequent to the time when normal
acceleration at the c.g. exceeds the positive limit maneuvering load
factor (at point A2 in Sec. 25.333(b)), or the resulting
tailplane normal load reaches its maximum, whichever occurs first, need
not be considered.
* * * * *
3. Section 25.335 is amended by revising paragraphs (a)(2) and
(b)(2) to read as follows:
Sec. 25.335 Design airspeeds.
* * * * *
(a) * * *
(2) Except as provided in Sec. 25.335(d)(2), VC may not
be less than VB + 1.32 U REF (with
UREF as specified in Sec. 25.341(a)(5)(i)). However
VC need not exceed the maximum speed in level flight at
maximum continuous power for the corresponding altitude.
* * * * *
(b) * * *
(2) The minimum speed margin must be enough to provide for
atmospheric variations (such as horizontal gusts, and penetration of
jet streams and cold fronts) and for instrument errors and airframe
production variations. These factors may be considered on a probability
basis. The margin at altitude where MC is limited by
compressibility effects must not less than 0.07M unless a lower margin
is determined using a rational analysis that includes the effects of
any automatic systems. In any case, the margin may not be reduced to
less than 0.05M.
* * * * *
4. Section 25.345 is amended by revising paragraph (d) to read as
follows:
Sec. 25.345 High lift devices.
* * * * *
(d) The airplane must be designed for a maneuvering load factor of
1.5 g at the maximum take-off weight with the wing-flaps and similar
high lift devices in the landing configurations.
5. Section 25.351 is revised to read as follows:
Sec. 25.351 Yaw maneuver conditions.
The airplane must be designed for loads resulting from the yaw
maneuver conditions specified in paragraphs (a) through (d) of this
section at speeds from VMC to VD. Unbalanced
aerodynamic moments about the center of gravity must be reacted in a
rational or conservative manner considering the airplane inertia
forces. In computing the tail loads the yawing velocity may be assumed
to be zero.
(a) With the airplane in unaccelerated flight at zero yaw, it is
assumed that the cockpit rudder control is suddenly displaced to
achieve the resulting rudder deflection, as limited by:
(1) The control system on control surface stops; or
(2) A limit pilot force of 300 pounds from VMC to
VA and 200 pounds from VC/MC to
VD/MD, with a linear variation between
VA and VC/MC.
(b) With the cockpit rudder control deflected so as always to
maintain the maximum rudder deflection available within the limitations
specified in paragraph (a) of this section, it is assumed that the
airplane yaws to the overswing sideslip angle.
(c) With the airplane yawed to the static equilibrium sideslip
angle, it is assumed that the cockpit rudder control is held so as to
achieve the maximum rudder deflection available within the limitations
specified in paragraph (a) of this section.
(d) With the airplane yawed to the static equilibrium sideslip
angle of paragraph (c) of this section, it is assumed that the cockpit
rudder control is suddenly returned to neutral.
6. Section 25.363 is amended by revising the heading and paragraph
(a) to read as follows:
Sec. 25.363 Side load on engine and auxiliary power unit mounts.
(a) Each engine and auxiliary power unit mount and its supporting
structure must be designed for a limit load factor in lateral
direction, for the side load on the engine and auxiliary power unit
mount, at least equal to the maximum load factor obtained in the yawing
conditions but not less than--
(1) 1.33; or
(2) One-third of the limit load factor for flight condition A as
prescribed in Sec. 25.333(b).
* * * * *
7. Section 25.371 is revised to read as follows:
Sec. 25.371 Gyroscopic loads.
The structure supporting any engine or auxiliary power unit must be
designed for the loads including the gyroscopic loads arising from the
conditions specified in Secs. 25.331, 25.341(a), 25.349, 25.351,
25.473, 25.479, and 25.481, with the engine or auxiliary power unit at
the maximum rpm appropriate to the condition. For the purposes of
compliance with this section, the pitch maneuver in Sec. 25.331(c)(1)
must be carried out until the positive limit maneuvering load factor
(point A2 in Sec. 25.333(b)) is reached.
[[Page 40705]]
8. Section 25.415 is amended by revising paragraph (a)(2) to read
as follows:
Sec. 25.415 Ground gust conditions.
(a) * * *
(2) The control system stops nearest the surfaces, the control
system locks, and the parts of the systems (if any) between these stops
and locks and the control surface horns, must be designed for limit
hinge moments H, in foot pounds, obtained from the formula,
H=.0034KV2cS, where--
V=65 (wind speed in knots)
K=limit hinge moment factor for ground gusts derived in paragraph
(b) of this section.
c=mean chord of the control surface aft of the hinge line (ft);
S=area of the control surface aft of the hinge line (sq ft);
* * * * *
9. Section 25.473 is revised to read as follows:
Sec. 25.473 Landing load conditions and assumptions.
(a) For the landing conditions specified in Sec. 25.479 to
Sec. 25.485 the airplane is assumed to contact the ground--
(1) In the attitudes defined in Sec. 25.479 and Sec. 25.481;
(2) With a limit descent velocity of 10 fps at the design landing
weight (the maximum weight for landing conditions at maximum descent
velocity); and
(3) With a limit descent velocity of 6 fps at the design take-off
weight (the maximum weight for landing conditions at a reduced descent
velocity).
(4) The prescribed descent velocities may be modified if it is
shown that the airplane has design features that make it impossible to
develop these velocities.
(b) Airplane lift, not exceeding airplane weight, may be assumed
unless the presence of systems or procedures significantly affects the
lift.
(c) The method of analysis of airplane and landing gear loads must
take into account at least the following elements:
(1) Landing gear dynamic characteristics.
(2) Spin-up and springback.
(3) Rigid body response.
(4) Structural dynamic response of the airframe, if significant.
(d) The limit inertia load factors corresponding to the required
limit descent velocities must be validated by tests as defined in
Sec. 25.723(a)
(e) The coefficient of friction between the tires and the ground
may be established by considering the effects of skidding velocity and
tire pressure. However, this coefficient of friction need not be more
than 0.8.
10. Section 25.479 is revised to read as follows:
Sec. 25.479 Level landing conditions.
(a) In the level attitude, the airplane is assumed to contact the
ground at forward velocity components, ranging from VL1 to
1.25 VL2 parallel to the ground under the conditions
prescribed in Sec. 25.473 with--
(1) VL1 equal to VS0 (TAS) at the appropriate
landing weight and in standard sea level conditions; and
(2) VL2 equal to VS0 (TAS) at the appropriate
landing weight and altitudes in a hot day temperature of 41 degrees F.
above standard.
(3) The effects of increased contact speed must be investigated if
approval of downwind landings exceeding 10 knots is requested.
(b) For the level landing attitude for airplanes with tail wheels,
the conditions specified in this section must be investigated with the
airplane horizontal reference line horizontal in accordance with Figure
2 of Appendix A of this part.
(c) For the level landing attitude for airplanes with nose wheels,
shown in Figure 2 of Appendix A of this part, the conditions specified
in this section must be investigated assuming the following attitudes:
(1) An attitude in which the main wheels are assumed to contact the
ground with the nose wheel just clear of the ground; and
(2) If reasonably attainable at the specified descent and forward
velocities, an attitude in which the nose and main wheels are assumed
to contact the ground simultaneously.
(d) In addition to the loading conditions prescribed in paragraph
(a) of this section, but with maximum vertical ground reactions
calculated from paragraph (a), the following apply:
(1) The landing gear and directly affected attaching structure must
be designed for the maximum vertical ground reaction combined with an
aft acting drag component of not less than 25% of this maximum vertical
ground reaction.
(2) The most severe combination of loads that are likely to arise
during a lateral drift landing must be taken into account. In absence
of a more rational analysis of this condition, the following must be
investigated:
(i) A vertical load equal to 75% of the maximum ground reaction of
Sec. 25.473 must be considered in combination with a drag and side load
of 40% and 35% respectively of that vertical load.
(ii) The shock absorber and tire deflections must be assumed to be
75% of the deflection corresponding to the maximum ground reaction of
Sec. 25.473(a)(2). This load case need not be considered in combination
with flat tires.
(3) The combination of vertical and drag components is considered
to be acting at the wheel axle centerline.
11. Section 25.481 is amended by revising paragraph (a)
introductory text and by designating the undesignated text following
paragraph (a)(2) as paragraph (a)(3) and revising it to read as
follows:
Sec. 25.481 Tail down landing conditions.
(a) In the tail-down attitude, the airplane is assumed to contact
the ground at forward velocity components, ranging from VL1
to VL2 parallel to the ground under the conditions
prescribed in Sec. 25.473 with--
(1) * * *
(2) * * *
(3) The combination of vertical and drag components is considered
to be acting at the main wheel axle centerline.
* * * * *
12. Section 25.483 is amended by revising the heading, introductory
text, and paragraph (a) to read as follows:
Sec. 25.483 One-gear landing conditions.
For the one-gear landing conditions, the airplane is assumed to be
in the level attitude and to contact the ground on one main landing
gear, in accordance with Figure 4 of Appendix A of this part. In this
attitude--
(a) The ground reactions must be the same as those obtained on that
side under Sec. 25.479(d)(1), and
* * * * *
13. Section 25.485 is amended by adding the introductory text to
read as follows:
Sec. 25.485 Side load conditions.
In addition to Sec. 25.479(d)(2) the following conditions must be
considered:
* * * * *
14. Section 25.491 is revised to read as follows:
Sec. 25.491 Taxi, takeoff and landing roll.
Within the range of appropriate ground speeds and approved weights,
the airplane structure and landing gear are assumed to be subjected to
loads not less than those obtained when the aircraft is operating over
the roughest ground that may reasonably be expected in normal
operation.
15. Section 25.499 is amended by revising the heading and paragraph
(e) to read as follows:
Sec. 25.499 Nose-wheel yaw and steering.
* * * * *
[[Page 40706]]
(e) With the airplane at design ramp weight, and the nose gear in
any steerable position, the combined application of full normal
steering torque and vertical force equal to 1.33 times the maximum
static reaction on the nose gear must be considered in designing the
nose gear, its attaching structure, and the forward fuselage structure.
16. Section 25.561 is amended by revising paragraph (c) to read as
follows:
Sec. 25.561 General.
* * * * *
(c) For equipment, cargo in the passenger compartments and any
other large masses, the following apply:
(1) Except as provided in paragraph (c)(2) of this section, these
items must be positioned so that if they break loose they will be
unlikely to:
(i) Cause direct injury to occupants;
(ii) Penetrate fuel tanks or lines or cause fire or explosion
hazard by damage to adjacent systems; or
(iii) Nullify any of the escape facilities provided for use after
an emergency landing.
(2) When such positioning is not practical (e.g. fuselage mounted
engines or auxiliary power units) each such item of mass shall be
restrained under all loads up to those specified in paragraph (b)(3) of
this section. The local attachments for these items should be designed
to withstand 1.33 times the specified loads if these items are subject
to severe wear and tear through frequent removal (e.g. quick change
interior items).
* * * * *
Issued in Washington D.C. on July 14, 1997.
Barry L. Valentine,
Acting Administrator.
[FR Doc. 97-19040 Filed 7-28-97; 8:45 am]
BILLING CODE 4910-13-M