Special Conditions: Bell Helicopter Textron, Inc. (BHTI), Model 525 Helicopters; Interaction of Systems and Structures., 88616-88619 [2016-29431]

Agencies

• Federal Aviation Administration
• DEPARTMENT OF TRANSPORTATION

[Federal Register Volume 81, Number 236 (Thursday, December 8, 2016)]
[Rules and Regulations]
[Pages 88616-88619]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2016-29431]



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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 29

[Docket No. FAA-2016-6939; Notice No. 29-038-SC]


Special Conditions: Bell Helicopter Textron, Inc. (BHTI), Model 
525 Helicopters; 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 BHTI Model 525 
helicopter. This helicopter will have a novel or unusual design feature 
associated with fly-by-wire flight control system (FBW FCS) functions 
that affect the structural integrity of the rotorcraft. The applicable 
airworthiness regulations do not contain adequate or appropriate safety 
standards for this design feature. These special conditions contain the 
additional safety standards that 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 9, 2017.

FOR FURTHER INFORMATION CONTACT: Martin R. Crane, Aviation Safety 
Engineer, Safety Management Group, Rotorcraft Directorate, FAA, 10101 
Hillwood Pkwy, Fort Worth, TX 76177; telephone (817) 222-5110; email 
martin.r.crane@faa.gov.

SUPPLEMENTARY INFORMATION: 

Background

    On December 15, 2011, BHTI applied for a type certificate for a new 
transport category helicopter designated as the Model 525. The aircraft 
is a medium twin engine rotorcraft. The design maximum takeoff weight 
is 20,000 pounds, with a maximum capacity of 16 passengers and a crew 
of 2.
    The BHTI Model 525 helicopter will be equipped with a FBW FCS. The 
control functions of the FBW FCS and its related systems affect the 
structural integrity of the rotorcraft. Current regulations do not take 
into account loads for the rotorcraft due to the effects of systems on 
structural performance including normal operation and failure 
conditions with strength levels related to probability of occurrence. 
Special conditions are needed to account for these features.

Type Certification Basis

    Under the provisions of 14 CFR 21.17, BHTI must show that the Model 
525 helicopter meets the applicable provisions of part 29, as amended 
by Amendment 29-1 through 29-55 thereto. The BHTI Model 525 
certification basis date is December 15, 2011, the date of application 
to the FAA.
    If the Administrator finds that the applicable airworthiness 
regulations (i.e., 14 CFR part 29) do not contain adequate or 
appropriate safety standards for the BHTI Model 525 because of a novel 
or unusual design feature, special conditions are prescribed under the 
provisions of Sec.  21.16.
    Special conditions are initially applicable to the model for which 
they are issued. Should the type certificate for that model be amended 
later to include any other model that incorporates the same or similar 
novel or unusual design feature, the special conditions would also 
apply to the other model under Sec.  21.101.
    In addition to the applicable airworthiness regulations and special 
conditions, the BHTI Model 525 helicopter must comply with the noise 
certification requirements of 14 CFR part 36, and the FAA must issue a 
finding of regulatory adequacy under Sec.  611 of Public Law 92-574, 
the ``Noise Control Act of 1972.''
    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.  21.17(a)(2).

Novel or Unusual Design Features

    The BHTI Model 525 helicopter will incorporate the following novel 
or unusual design features: FBW FCS, and its related systems (stability 
augmentation system, load alleviation system, flutter control system, 
and fuel management system), with control functions that affect the 
structural integrity of the rotorcraft. Current regulations are 
inadequate for considering the effects of these systems and their 
failures on structural performance. The general approach of accounting 
for the effect of system failures on structural performance would be 
extended to include any system where partial or complete failure, alone 
or in combination with any other system's partial or complete failure, 
would affect structural performance.

Discussion

    Active flight control systems are capable of providing automatic 
responses to inputs from sources other than the pilots. Active flight 
control systems have been expanded in function, effectiveness, and 
reliability to the point that FBW FCS systems are being installed on 
new rotorcraft. As a result of these advancements in flight control 
technology, 14 CFR part 29 does not provide a basis to achieve an 
acceptable level of safety for rotorcraft so equipped. Certification of 
these systems requires issuing special conditions under the provisions 
of Sec.  21.16.
    In the past, traditional rotorcraft flight control system designs 
have incorporated power-operated systems, stability or control 
augmentation with limited control authority, and autopilots that were 
certificated partly under Sec.  29.672 with guidance from Advisory 
Circular 29-2C, Section AC 29.672. These systems are integrated into 
the primary flight controls and are given sufficient control authority 
to maneuver the rotorcraft up to its structural design limits in 14 CFR 
part 29 subparts C and D. The FBW FCS advanced technology with its full 
authority necessitates additional requirements to account for the 
interaction of control systems and structures.
    The regulations defining the loads envelope in 14 CFR part 29 do 
not fully account for the effects of systems on structural performance. 
Automatic systems may be inoperative or they may operate in a degraded 
mode with less than full system authority and associated built-in 
protection features. Therefore, it is necessary to determine the 
structural factors of safety and operating margins such that the 
probability of structural failures due to application of loads during 
FBW FCS malfunctions is not greater than that found in rotorcraft 
equipped with traditional flight control systems. To achieve this 
objective and to ensure an acceptable level of safety, it is necessary 
to define the failure conditions and their associated frequency of 
occurrence.
    Traditional flight control systems provide two states, either fully 
functioning or completely inoperative. These conditions are readily 
apparent to the flight crew. Newer active flight control systems have 
failure modes that allow the system to function in a degraded mode 
without full authority and associated built-in protection features. As 
these degraded modes are not readily apparent to the flight crew, 
monitoring systems are required to provide an annunciation of degraded 
system capability.

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Comments

    A notice of proposed special conditions for the BHTI Model 525 
helicopter FBW FCS and its related systems was published in the Federal 
Register on May 27, 2016 (81 FR 33606). We did not receive any 
comments.

Applicability

    As discussed above, these special conditions are applicable to the 
BHTI Model 525 helicopter. Should BHTI apply at a later date for a 
change to the type certificate to include another model incorporating 
the same novel or unusual design feature, the special conditions would 
apply to that model as well.

Conclusion

    This action affects only certain novel or unusual design features 
on one model of rotorcraft. It is not a rule of general applicability.

List of Subjects in 14 CFR Part 29

    Aircraft, Aviation safety, Reporting and recordkeeping 
requirements.

    The authority citation for these special conditions is as follows:

    Authority: 49 U.S.C. 106(g), 40113, 44701, 44702, 44704.

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 Bell Helicopter Textron, Inc., Model 
525 helicopters when a fly-by-wire flight control system is installed:

Interaction of Systems and Structures

    For rotorcraft equipped with systems that affect structural 
performance, either directly or as a result of a failure or 
malfunction, the influence of these systems and their failure 
conditions must be taken into account when showing compliance with the 
requirements of Title 14, Code of Federal Regulations (14 CFR) part 29 
subparts C and D.
    The following criteria must be used for showing compliance with 
these special conditions for rotorcraft 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.
    (a) The criteria defined herein only address the direct structural 
consequences of the system responses and performance. They cannot be 
considered in isolation but should be included in the overall safety 
evaluation of the rotorcraft. 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 these special conditions.
    (b) Depending upon the specific characteristics of the rotorcraft, 
additional studies may be required that go beyond the criteria provided 
in this special condition in order to demonstrate the capability of the 
rotorcraft to meet other realistic conditions such as alternative gust 
or maneuver descriptions for a rotorcraft equipped with a load 
alleviation system.
    (c) The following definitions are applicable to these special 
conditions:
    (1) Structural performance: Capability of the rotorcraft to meet 
the structural requirements of 14 CFR part 29.
    (2) Flight limitations: Limitations that can be applied to the 
rotorcraft flight conditions following an in-flight occurrence and that 
are included in the flight manual (e.g., speed limitations and 
avoidance of severe weather conditions).
    (3) Operational limitations: Limitations, including flight 
limitations, which can be applied to the rotorcraft operating 
conditions before dispatch (e.g., fuel, payload, and Master Minimum 
Equipment List limitations).
    (4) Probabilistic terms: The terms ``improbable'' and ``extremely 
improbable'' are the same as those used in Sec.  29.1309.
    (5) Failure condition: The term ``failure condition'' is the same 
as that used in Sec.  29.1309; however, these special conditions apply 
only to system failure conditions that affect the structural 
performance of the rotorcraft (e.g., system failure conditions that 
induce loads, change the response of the rotorcraft to inputs such as 
gusts or pilot actions, or lower flutter margins).

Effects of Systems on Structures

    (a) General. The following criteria will be used in determining the 
influence of a system and its failure conditions on the rotorcraft 
structure.
    (b) 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 rotorcraft 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 rotorcraft must meet the strength requirements of part 29 
(static strength, residual strength), 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 rotorcraft has 
design features that will not allow it to exceed those limit 
conditions.
    (3) The rotorcraft must meet the flutter and divergence 
requirements of Sec.  29.629.
    (c) System in the failure condition. For all system failure 
conditions shown to be not 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 the 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 the ultimate loads that must be 
considered for design. The factor of safety is defined in Figure 1.

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[GRAPHIC] [TIFF OMITTED] TR08DE16.007

    (ii) For residual strength substantiation, the rotorcraft must be 
able to withstand two-thirds of the ultimate loads defined in paragraph 
(c)(1)(i) of these special conditions.
    (iii) Freedom from flutter and divergence must be shown under all 
conditions of operation including:
    (A) Airspeeds up to 1.11 VNE (power on and power off).
    (B) Main rotor speeds from 0.95 multiplied by the minimum permitted 
speed up to 1.05 multiplied by the maximum permitted speed (power on 
and power off).
    (C) The critical combinations of weight, center of gravity 
position, load factor, and altitude.
    (iv) For failure conditions that result in excursions beyond 
operating limitations, freedom from flutter and divergence must be 
shown to increased speeds, so that the margins intended by paragraph 
(c)(1)(iii) of these special conditions 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 rotorcraft in the 
system failed state, and considering all appropriate reconfiguration 
and flight limitations, the following apply:
    (i) The loads derived from the following conditions (or defined by 
special conditions or equivalent level of safety in lieu of the 
following conditions) at speeds up to VNE (power on and 
power off) (or the speed limitation prescribed for the remainder of the 
flight) and at the minimum and maximum main rotor speeds, if 
applicable, must be determined:
    (A) The limit maneuvering conditions specified in Sec. Sec.  29.337 
and 29.339.
    (B) The limit gust conditions specified in Sec.  29.341.
    (C) The limit yaw maneuvering conditions specified in Sec.  29.351.
    (D) The limit unsymmetrical conditions specified in Sec.  29.427.
    (E) The limit ground loading conditions specified in Sec.  29.473.
    (ii) For static strength substantiation, each part of the structure 
must be able to withstand the loads in paragraph (c)(2)(i) of these 
special conditions 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] TR08DE16.008

Qj = (Tj)(Pj)

Where:

Tj = Average time spent in failure condition j (in hours)
Pj = Probability of occurrence of failure mode j (per 
hour)


    Note: If Pj is greater than 10-\3\ per 
flight hour, then a 1.5 factor of safety must be applied to all 
limit load conditions specified in Subpart C.

    (iii) For residual strength substantiation, the rotorcraft must be 
able to withstand two-thirds of the ultimate loads defined in paragraph 
(c)(2)(ii) of these special conditions.
    (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 flutter and divergence must be shown up to 1.11 
VNE (power on and power off).
    (vi) Freedom from flutter and divergence must also be shown up to 
1.11 VNE (power on and power off) for all probable system 
failure conditions combined with any damage required or considered 
under Sec.  29.571(g) or Sec.  29.573(d)(3).
    (3) Consideration of certain failure conditions may be required by 
other sections of 14 CFR part 29 regardless of calculated system 
reliability. Where the failure analysis shows the probability of

[[Page 88619]]

these failure conditions to be less than 10-\9\, criteria 
other than those specified in this paragraph may be used for structural 
substantiation to show continued safe flight and landing.
    (d) 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 14 CFR part 29 or that significantly reduce the 
reliability of the remaining operational portion of the system. As far 
as reasonably practicable, the flight crew 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 other means of detecting failures before flight will 
become part of the certification maintenance requirements (CMRs) and 
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, shown to be not 
extremely improbable, during flight that could significantly affect the 
structural capability of the rotorcraft and for which the associated 
reduction in airworthiness can be minimized by suitable flight 
limitations, must be signaled to the flight crew. For example, failure 
conditions that result in a factor of safety between the rotorcraft 
strength and the loads of Subpart C below 1.25, or flutter and 
divergence margins below 1.11 VNE (power on and power off), 
must be signaled to the crew during flight.
    (e) Dispatch with known failure conditions. If the rotorcraft is to 
be dispatched in a known system failure condition that affects 
structural performance, or that affects the reliability of the 
remaining operational portion of the system to maintain structural 
performance, then the provisions of these special conditions must be 
met, including the provisions of paragraph (b) for the dispatched 
condition and paragraph (c) for subsequent failures. Expected 
operational limitations may be taken into account in establishing 
Pj as the probability of failure occurrence for determining 
the safety margin in Figure 1 of these special conditions. Flight 
limitations and expected operational limitations may be taken into 
account in establishing Qj as the combined probability of 
being in the dispatched failure condition and the subsequent failure 
condition for the safety margins in Figure 2 of these special 
conditions. These limitations must be such that the probability of 
being in this combined failure state and then subsequently encountering 
limit load conditions is extremely improbable. No reduction in these 
safety margins is allowed if the subsequent system failure rate is 
greater than 10-\3\ per hour.

    Issued in Fort Worth, Texas, on November 30, 2012.
Lance Gant,
Manager, Rotorcraft Directorate, Aircraft Certification Service.
[FR Doc. 2016-29431 Filed 12-7-16; 8:45 am]
 BILLING CODE 4910-13-P