Special Conditions: Cranfield Aerospace Limited, Cessna Aircraft Company Model 525; Tamarack Load Alleviation System and Cranfield Winglets-Interaction of Systems and Structures, 1163-1169 [2016-31819]

Agencies

• Federal Aviation Administration
• DEPARTMENT OF TRANSPORTATION

[Federal Register Volume 82, Number 3 (Thursday, January 5, 2017)]
[Rules and Regulations]
[Pages 1163-1169]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2016-31819]



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Federal Register / Vol. 82, No. 3 / Thursday, January 5, 2017 / Rules 
and Regulations

[[Page 1163]]



DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 23

[Docket No.FAA-2016-9409; Special Conditions No. 23-279-SC]


Special Conditions: Cranfield Aerospace Limited, Cessna Aircraft 
Company Model 525; Tamarack Load Alleviation System and Cranfield 
Winglets--Interaction of Systems and Structures

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Final special conditions.

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SUMMARY: These special conditions are issued for the Cessna Aircraft 
Company model 525 airplane. This airplane as modified by Cranfield 
Aerospace Limited will have a novel or unusual design feature 
associated with the installation of a Tamarack Active Technology Load 
Alleviation System and Cranfield Winglets. The applicable airworthiness 
regulations do not contain adequate or appropriate safety standards for 
this design feature. These special conditions contain the additional 
safety standards the Administrator considers necessary to establish a 
level of safety equivalent to that established by the existing 
airworthiness standards.

DATES: These special conditions are effective January 5, 2017 and are 
applicable on December 23, 2016.

FOR FURTHER INFORMATION CONTACT: Mike Reyer, Continued Operational 
Safety, ACE-113, Small Airplane Directorate, Aircraft Certification 
Service, 901 Locust; Kansas City, Missouri 64106; telephone (816) 329-
4131; facsimile (816) 329-4090.

SUPPLEMENTARY INFORMATION:

Background

    On January 25, 2016, Cranfield Aerospace Limited (CAL) applied for 
a supplemental type certificate to install winglets on the Cessna 
Aircraft Company (Cessna) model 525. The Cessna model 525 twin turbofan 
engine airplane is certified in the normal category for eight seats, 
including a pilot, a maximum gross weight of 10,700 pounds, and a 
maximum altitude of 41,000 feet mean sea level.
    Special conditions have been applied on past 14 CFR part 25 
airplane programs in order to consider the effects of systems on 
structures. The regulatory authorities and industry developed 
standardized criteria in the Aviation Rulemaking Advisory Committee 
(ARAC) forum based on the criteria defined in Advisory Circular 25.672-
1, dated November 15, 1983. The ARAC recommendations have been 
incorporated in the European Aviation Safety Agency Certification 
Specifications (CS) 25.302 and CS 25, appendix K. The special 
conditions used for part 25 airplane programs, can be applied to part 
23 airplane programs in order to require consideration of the effects 
of systems on structures. However, some modifications to the part 25 
special conditions are necessary to address differences between parts 
23 and 25 as well as differences between parts 91 and 121 operating 
environments.
    Winglets increase aerodynamic efficiency. However, winglets also 
increase wing design static loads, increase the severity of the wing 
fatigue spectra, and alter the wing fatigue stress ratio, which under 
limit gust and maneuvering loads factors, may exceed the certificated 
wing design limits. The addition of the Tamarack Active Technology Load 
Alleviation System (ATLAS) mitigates the winglet's adverse structural 
effects by reducing the aerodynamic effectiveness of the winglet when 
ATLAS senses gust and maneuver loads above a predetermined threshold.
    The ATLAS functions as a load-relief system. This is accomplished 
by measuring airplane loading via an accelerometer and moving an 
aileron-like device called a Tamarack Active Control Surface (TACS) 
that reduces lift at the tip of the wing. The TACS are located outboard 
and adjacent to the left and right aileron control surfaces. The TACS 
movement reduces lift at the tip of the wing, resulting in the wing 
spanwise center of pressure moving inboard, thus reducing bending 
stresses along the wing span. Because the ATLAS compensates for the 
increased wing root bending at elevated load factors, the overall 
effect of this modification is that the required reinforcement of the 
existing Cessna wing structure due to the winglet installation is 
reduced. The applicable airworthiness regulations do not contain 
adequate or appropriate safety standards for this design feature.
    The ATLAS is not a primary flight control system, a trim device, or 
a wing flap. However, several regulations under Part 23, Subpart D--
Design and Construction--Control Systems, have applicability to ATLAS, 
which might otherwise be considered ``Not Applicable'' under a strict 
interpretation of the regulations. These Control System regulations 
include Sec. Sec.  23.672, 23.675, 23.677, 23.681, 23.683, 23.685, 
23.693, 23.697, and 23.701.
    An airplane designed with a load-relief system must provide an 
equivalent level of safety to an airplane with similar characteristics 
designed without a load-relief system. In the following special 
conditions, an equivalent level of safety is provided by relating the 
required structural safety factor to the probability of load-relief 
system failure and the probability of exceeding the frequency of design 
limit and ultimate loads.
    These special conditions address several issues with the operation 
and failure of the load-relief system. These issues include the 
structural requirements for the system in the fully operational state; 
evaluation of the effects of system failure, both at the moment of 
failure and continued safe flight and landing with the failure 
annunciated to the pilot; and the potential for failure of the failure 
monitoring/pilot annunciation function.
    The structural requirements for the load-relief system in the fully 
operational state are stated in special condition 2(e) of these special 
conditions. In this case, the structure must meet the full requirements 
of part 23, subparts C and D with full credit given for the effects of 
the load-relief system.
    In the event of a load-relief system failure in-flight, the effects 
on the structure at the moment of failure must be considered as 
described in special condition 2(f)(l) of these special

[[Page 1164]]

conditions. These effects include, but are not limited to the 
structural loads induced by a hard-over failure of the load-relief 
control surface and oscillatory system failures that may excite the 
structural dynamic modes. In evaluating these effects, pilot corrective 
actions may be considered and the airplane may be assumed to be in 1g 
(gravitation force) flight prior to the load-relief system failure. 
These special conditions allows credit, in the form of reduced 
structural factors of safety, based on the probability of failure of 
the load-relief system. Effects of an in-flight failure on flutter and 
fatigue and damage tolerance must also be evaluated.
    Following the initial in-flight failure, the airplane must be 
capable of continued safe flight and landing. Special condition 2(f)(2) 
in these special conditions assumes that a properly functioning, 
monitoring, and annunciating system has alerted the pilot to the load-
relief failure. Since the pilot has been made aware of the load-relief 
failure, appropriate flight limitations, including speed restrictions, 
may be considered when evaluating structural loads, flutter, and 
fatigue and damage tolerance. These special conditions allows credit, 
in the form of reduced structural factors of safety, based on the 
probability of failure of the load-relief system and the flight time 
remaining on the failure flight.
    Special condition 2(g) of these special conditions addresses the 
failure of the load-relief system to annunciate a failure to the pilot. 
These special conditions address this concern with maintenance actions 
and requirements for monitoring and annunciation systems.
    These special conditions have been modified from previous, similar 
part 25 special conditions because of the differences between parts 23 
and 25 as well as to address the part 91 operating and maintenance 
environment. Paragraph (c)(3) of the part 25 special condition \1\ is 
removed from these special conditions. Special condition 2(h) of these 
special conditions is modified to require a ferry permit for additional 
flights after an annunciated failure or obvious system failure.
---------------------------------------------------------------------------

    \1\ Special Condition No. 25-164-SC, ``Boeing Model 737-700 IGW, 
Interaction of Systems and Structures,'' Effective August 30, 2000 
(65 FR 55443).
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Type Certification Basis

    Under the provisions of Sec.  21.101, Cranfield Aerospace Limited 
must show that the Cessna model 525, as changed, continues to meet the 
applicable provisions of the regulations incorporated by reference in 
Type Certificate No. A1WI, revision 24, or the applicable regulations 
in effect on the date of application for the change. The regulations 
incorporated by reference in the type certificate are commonly referred 
to as the ``original type certification basis.'' The regulations 
incorporated by reference in Type Certificate No. A1WI, revision 24 are 
14 CFR part 23 effective February 1, 1965, amendments 23-1 through 23-
38 and 23-40.
    If the Administrator finds the applicable airworthiness regulations 
(i.e., 14 CFR part 23) do not contain adequate or appropriate safety 
standards for the Cessna model 525 because of a novel or unusual design 
feature, special conditions are prescribed under the provisions of 
Sec.  21.16.
    In addition to the applicable airworthiness regulations and special 
conditions, the Cessna 525 must comply with the fuel vent and exhaust 
emission requirements of 14 CFR part 34 and the noise certification 
requirements of 14 CFR part 36.
    The FAA issues special conditions, as defined in 14 CFR 11.19, in 
accordance with Sec.  11.38, and they become part of the type-
certification basis under Sec.  21.101.
    Special conditions are initially applicable to the model for which 
they are issued. Should the applicant apply for a supplemental type 
certificate to modify any other model included on the same type 
certificate to incorporate the same or similar novel or unusual design 
feature, the FAA would apply these special conditions to the other 
model under Sec.  21.101.

Novel or Unusual Design Features

    The Cessna model 525 will incorporate the following novel or 
unusual design features: Cranfield winglets with a Tamarack Active 
Technology Load Alleviation System.

Discussion

    For airplanes equipped with systems that affect structural 
performance, either directly or as a result of a failure or 
malfunction, the applicant must take into account the influence of 
these systems and their failure conditions when showing compliance with 
the requirements of part 23, subparts C and D.
    The applicant must use the following criteria for showing 
compliance with these special conditions for airplanes equipped with 
flight control systems, autopilots, stability augmentation systems, 
load alleviation systems, flutter control systems, fuel management 
systems, and other systems that either directly or as a result of 
failure or malfunction affect structural performance. If these special 
conditions are used for other systems, it may be necessary to adapt the 
criteria to the specific system.

Discussion of Comments

    Notice of proposed special conditions No. 23-16-03-SC for the 
Cessna model 525 airplane was published in the Federal Register on 
November 22, 2016 (81 FR 83737). No comments were received, and the 
special conditions are adopted as proposed.

Applicability

    As discussed above, these special conditions are applicable to the 
Cessna model 525. Should Cranfield Aerospace Limited apply at a later 
date for a supplemental type certificate to modify any other model 
included on A1WI, revision 24 to incorporate the same novel or unusual 
design feature, the FAA would apply these special conditions to that 
model as well.
    Under standard practice, the effective date of final special 
conditions would be 30 days after the date of publication in the 
Federal Register; however, as the supplemental type certification date 
for the Cessna model 525 is imminent, the FAA finds that good cause 
exists to make these special conditions effective upon issuance.

Conclusion

    This action affects only certain novel or unusual design features 
on one model of airplanes. It is not a rule of general applicability 
and it affects only the applicant who applied to the FAA for approval 
of these features on the airplane.

List of Subjects in 14 CFR Part 23

    Aircraft, Aviation safety, Signs and symbols.

    Authority:  49 U.S.C. 106(g), 40113, 44701, 44702, 14 CFR 21.16, 
21.101; and 14 CFR 11.38 and 11.19.

The Special Conditions

    Accordingly, pursuant to the authority delegated to me by the 
Administrator, the following special conditions are issued as part of 
the type certification basis for Cessna Aircraft Company 525 airplanes 
modified by Cranfield Aerospace Limited.

1. Active Technology Load Alleviation System (ATLAS)

SC 23.672 Load Alleviation System

    The load alleviation system must comply with the following:

[[Page 1165]]

    (a) A warning, which is clearly distinguishable to the pilot under 
expected flight conditions without requiring the pilot's attention, 
must be provided for any failure in the load alleviation system or in 
any other automatic system that could result in an unsafe condition if 
the pilot was not aware of the failure. Warning systems must not 
activate the control system.
    (b) The design of the load alleviation system or of any other 
automatic system must permit initial counteraction of failures without 
requiring exceptional pilot skill or strength, by either the 
deactivation of the system or a failed portion thereof, or by 
overriding the failure by movement of the flight controls in the normal 
sense.
    (1) If deactivation of the system is used to counteract failures, 
the control for this initial counteraction must be readily accessible 
to each pilot while operating the control wheel and thrust control 
levers.
    (2) If overriding the failure by movement of the flight controls is 
used, the override capability must be operationally demonstrated.
    (c) It must be shown that, after any single failure of the load 
alleviation system, the airplane must be safely controllable when the 
failure or malfunction occurs at any speed or altitude within the 
approved operating limitations that is critical for the type of failure 
being considered;
    (d) It must be shown that, while the system is active or after any 
single failure of the load alleviation system--
    (1) The controllability and maneuverability requirements of part 
23, subpart D, are met within a practical operational flight envelope 
(e.g., speed, altitude, normal acceleration, and airplane 
configuration) that is described in the Airplane Flight Manual (AFM); 
and
    (2) The trim, stability, and stall characteristics are not impaired 
below a level needed to permit continued safe flight and landing.

SC 23.677 Load Alleviation Active Control Surface

    (a) Proper precautions must be taken to prevent inadvertent or 
improper operation of the load alleviation system. It must be 
demonstrated that with the load alleviation system operating throughout 
its operational range, a pilot of average strength and skill level is 
able to continue safe flight with no objectionable increased workload.
    (b) The load alleviation system must be designed so that, when any 
one connecting or transmitting element in the primary flight control 
system fails, adequate control for safe flight and landing is 
available.
    (c) The load alleviation system must be irreversible unless the 
control surface is properly balanced and has no unsafe flutter 
characteristics. The system must have adequate rigidity and reliability 
in the portion of the system from the control surface to the attachment 
of the irreversible unit to the airplane structure.
    (d) It must be demonstrated the airplane is safely controllable and 
a pilot can perform all maneuvers and operations necessary to affect a 
safe landing following any load alleviation system runaway not shown to 
be extremely improbable, allowing for appropriate time delay after 
pilot recognition of the system runaway. The demonstration must be 
conducted at critical airplane weights and center of gravity positions.

SC 23.683 Operation Tests

    (a) It must be shown by operation tests that, when the flight 
control system and the load alleviation systems are operated and loaded 
as prescribed in paragraph (c) of this section, the flight control 
system and load alleviation systems are free from--
    (1) Jamming;
    (2) Excessive friction; and
    (3) Excessive deflection.
    (b) The operation tests in paragraph (a) of this section must also 
show the load alleviation system and associated surfaces do not 
restrict or prevent aileron control surface movements, or cause any 
adverse response of the ailerons, under the loading prescribed in 
paragraph (c) of this section that would prevent continued safe flight 
and landing.
    (c) The prescribed test loads are for the entire load alleviation 
and flight control systems, loads corresponding to the limit air loads 
on the appropriate surfaces.

    Note: Advisory Circular (AC) 23-17C ``Systems and Equipment 
Guide to Certification of Part 23 Airplanes'' provides guidance on 
potential methods of compliance with this section and other 
regulations applicable to this STC project.

SC 23.685 Control System Details

    (a) Each detail of the load alleviation system and related moveable 
surfaces must be designed and installed to prevent jamming, chafing, 
and interference from cargo, passengers, loose objects, or the freezing 
of moisture.
    (b) There must be means in the cockpit to prevent the entry of 
foreign objects into places where they would jam any one connecting or 
transmitting element of the load alleviation system.
    (c) Each element of the load alleviation system must have design 
features, or must be distinctively and permanently marked, to minimize 
the possibility of incorrect assembly that could result in 
malfunctioning of the control system.

SC 23.697 Load Alleviation System Controls

    (a) The load alleviation control surface must be designed so that 
during normal operation, when the surface has been placed in any 
position, it will not move from that position unless the control is 
adjusted or is moved by the operation of a load alleviation system.
    (b) The rate of movement of the control surface in response to the 
load alleviation system controls must give satisfactory flight and 
performance characteristics under steady or changing conditions of 
airspeed, engine power, attitude, flap configuration, speedbrake 
position, and during landing gear extension and retraction.

SC 23.701 Load Alleviation System Interconnection

    (a) The load alleviation system and related movable surfaces as a 
system must--
    (1) Be synchronized by a mechanical interconnection between the 
movable surfaces or by an approved equivalent means; or
    (2) Be designed so the occurrence of any failure of the system that 
would result in an unsafe flight characteristic of the airplane is 
extremely improbable; or
    (b) The airplane must be shown to have safe flight characteristics 
with any combination of extreme positions of individual movable 
surfaces.
    (c) If an interconnection is used in multiengine airplanes, it must 
be designed to account for unsymmetrical loads resulting from flight 
with the engines on one side of the plane of symmetry inoperative and 
the remaining engines at takeoff power. For single-engine airplanes, 
and multiengine airplanes with no slipstream effects on the load 
alleviation system, it may be assumed that 100 percent of the critical 
air load acts on one side and 70 percent on the other.

Sections 23.675, ``Stops;'' 23.681, ``Limit Load Static Tests;'' and 
23.693, ``Joints''

    The load alleviation system must comply with Sec. Sec.  23.675, 
23.681, and 23.693 as written and no unique special condition will be 
required for these regulations.

[[Page 1166]]

Applicability of Control System Regulations to Other Control Systems

    If applicable, other control systems used on the Cessna 525 may 
require a showing of compliance to Sec. Sec.  23.672, 23.675, 23.677, 
23.681, 23.683, 23.685, 23.693, 23.697 and 23.701 as written for this 
STC project.

2. Interaction of Systems and Structures

    (a) The criteria defined herein only address the direct structural 
consequences of the system responses and performances and cannot be 
considered in isolation but should be included in the overall safety 
evaluation of the airplane. These criteria may in some instances 
duplicate standards already established for this evaluation. These 
criteria are only applicable to structure whose failure could prevent 
continued safe flight and landing. Specific criteria that define 
acceptable limits on handling characteristics or stability requirements 
when operating in the system degraded or inoperative mode are not 
provided in this special condition.
    (b) Depending upon the specific characteristics of the airplane, 
additional studies may be required that go beyond the criteria provided 
in this special condition in order to demonstrate the capability of the 
airplane to meet other realistic conditions such as alternative gust or 
maneuver descriptions for an airplane equipped with a load alleviation 
system.
    (c) The following definitions are applicable to this special 
condition.
    (1) Structural performance: Capability of the airplane to meet the 
structural requirements of 14 CFR part 23.
    (2) Flight limitations: Limitations that can be applied to the 
airplane flight conditions following an in-flight occurrence and that 
are included in the flight manual (e.g., speed limitations, avoidance 
of severe weather conditions, etc.).
    (3) [Reserved]
    (4) Probabilistic terms: The probabilistic terms (probable, 
improbable, extremely improbable) used in this special condition are 
the same as those used in Sec.  23.1309. For the purposes of this 
special condition, extremely improbable for normal, utility, and 
acrobatic category airplanes is defined as 10-\8\ per hour. 
For commuter category airplanes, extremely improbable is defined as 
10-\9\ per hour.
    (5) Failure condition: The term failure condition is the same as 
that used in Sec.  23.1309, however this special condition applies only 
to system failure conditions that affect the structural performance of 
the airplane (e.g., system failure conditions that induce loads, change 
the response of the airplane to inputs such as gusts or pilot actions, 
or lower flutter margins).
    (d) General. The following criteria (paragraphs (e) through (i)) 
will be used in determining the influence of a system and its failure 
conditions on the airplane structure.
    (e) System fully operative. With the system fully operative, the 
following apply:
    (1) Limit loads must be derived in all normal operating 
configurations of the system from all the limit conditions specified in 
subpart C (or defined by special condition or equivalent level of 
safety in lieu of those specified in subpart C), taking into account 
any special behavior of such a system or associated functions or any 
effect on the structural performance of the airplane that may occur up 
to the limit loads. In particular, any significant nonlinearity (rate 
of displacement of control surface, thresholds or any other system 
nonlinearities) must be accounted for in a realistic or conservative 
way when deriving limit loads from limit conditions.
    (2) The airplane must meet the strength requirements of part 23 
(static strength and residual strength for failsafe or damage tolerant 
structure), using the specified factors to derive ultimate loads from 
the limit loads defined above. The effect of nonlinearities must be 
investigated beyond limit conditions to ensure the behavior of the 
system presents no anomaly compared to the behavior below limit 
conditions. However, conditions beyond limit conditions need not be 
considered when it can be shown that the airplane has design features 
that will not allow it to exceed those limit conditions.
    (3) The airplane must meet the aeroelastic stability requirements 
of Sec.  23.629.
    (f) System in the failure condition. For any system failure 
condition not shown to be extremely improbable, the following apply:
    (1) At the time of occurrence. Starting from 1-g level flight 
conditions, a realistic scenario, including pilot corrective actions, 
must be established to determine the loads occurring at the time of 
failure and immediately after failure.
    (i) For static strength substantiation, these loads, multiplied by 
an appropriate factor of safety that is related to the probability of 
occurrence of the failure, are ultimate loads to be considered for 
design. The factor of safety is defined in figure 1.

[[Page 1167]]

[GRAPHIC] [TIFF OMITTED] TR05JA17.316

    (ii) For residual strength substantiation, the airplane must be 
able to withstand two thirds of the ultimate loads defined in 
subparagraph (f)(1)(i).
    (iii) For pressurized cabins, these loads must be combined with the 
normal operating differential pressure.
    (iv) Freedom from aeroelastic instability must be shown up to the 
speeds defined in Sec.  23.629(f). For failure conditions that result 
in speeds beyond VD/MD, freedom from aeroelastic 
instability must be shown to increased speeds, so that the margins 
intended by Sec.  23.629(f) are maintained.
    (v) Failures of the system that result in forced structural 
vibrations (oscillatory failures) must not produce loads that could 
result in detrimental deformation of primary structure.
    (2) For the continuation of the flight. For the airplane, in the 
system failed state and considering any appropriate reconfiguration and 
flight limitations, the following apply:
    (i) The loads derived from the following conditions (or defined by 
special condition or equivalent level of safety in lieu of the 
following conditions) at speeds up to VC/MC, or 
the speed limitation prescribed for the remainder of the flight, must 
be determined:
    (A) The limit symmetrical maneuvering conditions specified in 
Sec. Sec.  23.321, 23.331, 23.333, 23.345, 23.421, 23.423, and 23.445.
    (B) The limit gust and turbulence conditions specified in 
Sec. Sec.  23.341, 23.345, 23.425, 23.443, and 23.445.
    (C) The limit rolling conditions specified in Sec.  23.349 and the 
limit unsymmetrical conditions specified in Sec. Sec.  23.347, 23.427, 
and 23.445.
    (D) The limit yaw maneuvering conditions specified in Sec. Sec.  
23.351, 23.441, and 23.445.
    (E) The limit ground loading conditions specified in Sec. Sec.  
23.473 and 23.493.
    (ii) For static strength substantiation, each part of the structure 
must be able to withstand the loads in paragraph (f)(2)(i) of this 
special condition multiplied by a factor of safety depending on the 
probability of being in this failure state. The factor of safety is 
defined in figure 2.

[[Page 1168]]

[GRAPHIC] [TIFF OMITTED] TR05JA17.317

    (iii) For residual strength substantiation, the airplane must be 
able to withstand two thirds of the ultimate loads defined in paragraph 
(f)(2)(ii) of this special condition. For pressurized cabins, these 
loads must be combined with the normal operating pressure differential.
    (iv) If the loads induced by the failure condition have a 
significant effect on fatigue or damage tolerance then their effects 
must be taken into account.
    (v) Freedom from aeroelastic instability must be shown up to a 
speed determined from figure 3. Flutter clearance speeds V' and V'' may 
be based on the speed limitation specified for the remainder of the 
flight using the margins defined by Sec.  23.629.

[[Page 1169]]

[GRAPHIC] [TIFF OMITTED] TR05JA17.318

    (vi) Freedom from aeroelastic instability must also be shown up to 
V' in figure 3 above, for any probable system failure condition 
combined with any damage required or selected for investigation by 
Sec. Sec.  23.571 through 23.574.
    (3) Consideration of certain failure conditions may be required by 
other sections of 14 CFR part 23 regardless of calculated system 
reliability. Where analysis shows the probability of these failure 
conditions to be less than 10-\8\ for normal, utility, or 
acrobatic category airplanes or less than 10-\9\ for 
commuter category airplanes, criteria other than those specified in 
this paragraph may be used for structural substantiation to show 
continued safe flight and landing.
    (g) Failure indications. For system failure detection and 
indication, the following apply:
    (1) The system must be checked for failure conditions, not 
extremely improbable, that degrade the structural capability below the 
level required by part 23 or significantly reduce the reliability of 
the remaining system. As far as reasonably practicable, the flightcrew 
must be made aware of these failures before flight. Certain elements of 
the control system, such as mechanical and hydraulic components, may 
use special periodic inspections, and electronic components may use 
daily checks, in lieu of detection and indication systems to achieve 
the objective of this requirement. These certification maintenance 
requirements must be limited to components that are not readily 
detectable by normal detection and indication systems and where service 
history shows that inspections will provide an adequate level of 
safety.
    (2) The existence of any failure condition, not extremely 
improbable, during flight that could significantly affect the 
structural capability of the airplane and for which the associated 
reduction in airworthiness can be minimized by suitable flight 
limitations, must be signaled to the flightcrew. The probability of not 
annunciating these failure conditions must be extremely improbable 
(unannunciated failure). For example, failure conditions that result in 
a factor of safety between the airplane strength and the loads of 
subpart C below 1.25, or flutter margins below V'', must be signaled to 
the flightcrew during flight.
    (h) Further flights with known load-relief system failure. 
Additional flights after an annunciated failure of the load-relief 
system or obvious failure of the load-relief system are permitted with 
a ferry permit only. In these cases, ferry permits may be issued to 
allow moving the airplane to an appropriate maintenance facility. 
Additional flights are defined as, further flights after landing on a 
flight where an annunciated or obvious failure of the load-relief 
system has occurred or after an annunciated or obvious failure of the 
load-relief system occurs during preflight preparation.
    (i) Fatigue and damage tolerance. If any system failure would have 
a significant effect on the fatigue or damage evaluations required in 
Sec. Sec.  23.571 through 23.574, then these effects must be taken into 
account.

    Issued in Kansas City, Missouri, on December 23, 2016.
Barry Ballenger,
Acting Manager, Small Airplane Directorate, Aircraft Certification 
Service.
[FR Doc. 2016-31819 Filed 1-4-17; 8:45 am]
 BILLING CODE 4910-13-P