[Federal Register Volume 82, Number 3 (Thursday, January 5, 2017)]
[Rules and Regulations]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2016-31819]
Rules and Regulations
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Federal Register / Vol. 82, No. 3 / Thursday, January 5, 2017 / Rules
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.
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
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.
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
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
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).
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
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.
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.
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.
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:
(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
(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
(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
(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);
(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
(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--
(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
(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
(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
(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
(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
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.
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
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
(c) The following definitions are applicable to this special
(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.).
(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
(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.
[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
(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
(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,
(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.
[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.
[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
(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
Issued in Kansas City, Missouri, on December 23, 2016.
Acting Manager, Small Airplane Directorate, Aircraft Certification
[FR Doc. 2016-31819 Filed 1-4-17; 8:45 am]
BILLING CODE 4910-13-P