[Federal Register Volume 88, Number 195 (Wednesday, October 11, 2023)]
[Rules and Regulations]
[Pages 70344-70351]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2023-22492]


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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 21

[Docket No. FAA-2023-0623]


Policy for Type Certification of Very Light Airplanes as a 
Special Class of Aircraft

AGENCY: Federal Aviation Administration (FAA), Department of 
Transportation (DOT).

ACTION: Notification of policy.

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SUMMARY: The FAA announces the policy for the type certification of 
Very Light Airplanes (VLA) as a special class of aircraft under the 
Federal Aviation Regulations.

DATES: This policy is effective October 11, 2023.

FOR FURTHER INFORMATION CONTACT: Hieu Nguyen, Product Policy 
Management, AIR-62B, Policy and Standards Division, Aircraft 
Certification Service, Federal Aviation Administration; telephone 816-
329-4123; email <a href="/cdn-cgi/l/email-protection#036b6a66762d6d64767a666d436562622d646c75"><span class="__cf_email__" data-cfemail="3e56575b4b1050594b475b507e585f5f10595148">[email protected]</span></a>.

SUPPLEMENTARY INFORMATION:

Background

    The FAA issued a notice of proposed policy, which published in the 
Federal Register on August 9, 2023 (88 FR 53815). The FAA received 
comments from two commenters. The comments are available to view in 
Docket No. FAA-2023-0623 at www.regulations.gov.

Discussion of Comments

    The FAA received one comment from an individual that was unrelated 
to the notice and outside the scope of the proposed policy. The other 
comment was a request from the General Aviation Manufacturers 
Association asking for a 30-day extension to the comment period. 
However, the FAA did not extend the comment period. The FAA chose a 30-
day comment period because it balances the need to have a final policy 
available for applicants with the need for interested persons to have 
time to comment on the proposed policy. The FAA determined that a 30-
day comment period provided adequate time for interested persons to 
submit comments and that it would not be in the public interest to 
extend the comment period.

Authority Citation

    The authority citations for these airworthiness criteria are as 
follows:

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

Policy

    The FAA will continue to allow type certification of VLA as a 
special class of aircraft under 14 CFR 21.17(b) using CS-VLA or JAR-VLA 
requirements, while also allowing eligibility for certification as a 
normal category airplane in accordance with part 23 using accepted 
means of compliance. The FAA accepts CS-VLA and JAR-VLA airworthiness 
criteria as providing an equivalent level of safety under Sec.  
21.17(b) special class type certification of VLA airplanes. The FAA 
will consider proposals for airplane designs that differ from the VLA 
limits defined in AC 21.17-3 for type certification as a special class 
of aircraft under Sec.  21.17(b), provided the VLA were certificated to 
the JAR-VLA or CS-VLA requirements plus additional airworthiness 
criteria the FAA finds appropriate and applicable for the proposed 
design. Additional design requirements may include but are not limited 
to the airworthiness criteria identified in the following paragraphs. 
Other additional airworthiness criteria may be required to address 
specific design proposals.

Advanced Avionic Displays

    If the airplane has advanced avionic displays installed, the 
following requirements from 14 CFR part 23 apply:
    <bullet> 14 CFR 23.1307 at amendment 23-49, Miscellaneous 
Equipment.
    <bullet> 14 CFR 23.1311 at amendment 23-62, Electronic Display 
Instrument Systems.
    <bullet> 14 CFR 23.1321 at amendment 23-49, Arrangement and 
Visibility.
    <bullet> 14 CFR 23.1359 at amendment 23-49, Electrical System Fire 
Protection.

Winglets

    If the airplane has any outboard fins or winglets installed, the 
design must comply with JAR 23.445.

Engine Mount to Composite Airframe

VLA.001
    The requirements in this section are applicable to airplanes with 
an engine mounting to composite airframe. Tests must be performed that 
demonstrate that the interface between the metallic engine mount and 
the glass fiber reinforced plastic fuselage withstand a fire for 15 
minutes while carrying loads under the following conditions:
    (a) With one lost engine mount fitting the loads are distributed 
over the remaining three engine mount fittings. The most critical of 
these fittings must be chosen for the test.
    The loads are:
    (1) In Z-direction the mass of the propulsion unit multiplied by a 
maneuvering load factor resulting from a 30[deg] turn for 15 minutes, 
superimposed by a maneuvering load of 3 seconds representing the 
maximum positive limit maneuvering load factor of n=3.8 from JAR-VLA 
337(a).
    (2) In X-direction the engine propulsion force at maximum 
continuous power for 5 minutes.
    (b) The flame to which the component test arrangement is subjected 
must provide a temperature of 500 [deg]C within the target area.
    (c) The flame must be large enough to maintain the required 
temperature over the entire test zone, i.e., the fitting on the engine 
compartment side.
    (d) It must be shown that the test equipment, e.g., burner and 
instrumentation are of sufficient power, size, and precision to yield 
the test requirements arising from paragraphs (a) through (c) of this 
section.

Night-VFR Operations

VLA.005
    The requirements in sections VLA.005 through VLA.105 are applicable 
to airplanes with a single engine (spark- or compression-ignition) 
having not more than two seats, with a maximum certificated takeoff 
weight of not more than 750 kg and a stalling speed in the landing 
configuration of not more than 83 km/h (45 knots)(CAS), to be approved 
for day-VFR [visual flight rules] or for day-and night-VFR.
VLA.010
    (a) Any short period oscillation not including combined lateral-
directional oscillations occurring between the stalling speed and the 
maximum allowable speed appropriate to the

[[Page 70345]]

configuration of the airplane must be heavily damped with the primary 
controls--
    (1) Free; and
    (2) In a fixed position.
    (b) Any combined lateral-directional oscillations (``Dutch roll'') 
occurring between the stalling speed and the maximum allowable speed 
appropriate to the configuration of the airplane must be damped to 1/10 
amplitude in 7 cycles with the primary controls--
    (1) Free; and
    (2) In a fixed position.
    (c) Any long period oscillation of the flight path (phugoid) must 
not be so unstable as to cause an unacceptable increase in pilot 
workload or otherwise endanger the airplane. When under the conditions 
specified in CS-VLA 175, the longitudinal control force required to 
maintain speeds differing from the trimmed speed by at least plus or 
minus 15% is suddenly released, the response of the airplane must not 
exhibit any dangerous characteristics nor be excessive in relation to 
the magnitude of the control force released.
VLA.015
    The pilot compartment must be free from glare and reflections that 
could interfere with the pilot's vision under all operations for which 
the certification is requested. The pilot compartment must be designed 
so that--
    (a) The pilot's view is sufficiently extensive, clear, and 
undistorted, for safe operation;
    (b) The pilot is protected from the elements so that moderate rain 
conditions do not unduly impair the pilot's view of the flight path in 
normal flight and while landing; and
    (c) Internal fogging of the windows covered under paragraph (a) of 
this section can be easily cleared by the pilot unless means are 
provided to prevent fogging.
VLA.020
    (a) The airplane must be so designed that unimpeded and rapid 
escape is possible in any normal and crash attitude.
    (b) The opening system must be designed for simple and easy 
operation. It must function rapidly and be designed so that it can be 
operated by each occupant strapped in their seat, and also from outside 
the cockpit. Reasonable provisions must be provided to prevent jamming 
by fuselage deformation.
    (c) The exit must be marked for easy location and operation even in 
darkness.
VLA.025
    (a) The engine must meet the specifications of CS-E, amendment 
6,\1\ or 14 CFR part 33, amendment 33-36, for night-VFR operation.
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    \1\ CS-E amendment 6: Certification Specifications and 
Acceptable Means of Compliance for Engines can be found in Docket 
No. FAA-2023-0623 at https://www.regulations.gov.
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    (b) Restart capability. An altitude and airspeed envelope must be 
established for the airplane for in-flight engine restarting and the 
installed engine must have a restart capability within that envelope.
VLA.030
    (a) For day-VFR operation, the propeller must meet the 
specifications of CS-22 Subpart J, amendment 3. For night-VFR 
operations the propeller and its control system must meet the 
specifications of CS-P, amendment 2,\2\ or 14 CFR part 35, amendment 
35-10, except for fixed pitch propellers, for which CS-22 \3\ subpart J 
is sufficient.
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    \2\ CS-P amendment 2: Certification Specifications and 
Acceptable Means of Compliance for Propellers can be found in Docket 
FAA-2023-0623 at https://www.regulations.gov.
    \3\ CS-22 amendment 3: Certification Specifications, Acceptable 
Means of Compliance and Guidance Material for Sailplanes and Powered 
Sailplanes can be found in Docket No. FAA-2023-0623 at https://www.regulations.gov.
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    (b) Engine power and propeller shaft rotational speed may not 
exceed the limits for which the propeller is certificated or approved.
VLA.035
    If an air filter is used to protect the engine against foreign 
material particles in the induction air supply--
    (a) Each air filter must be capable of withstanding the effects of 
temperature extremes, rain, fuel, oil, and solvents to which it is 
expected to be exposed in service and maintenance; and
    (b) Each air filter must have a design feature to prevent material 
separated from the filter media from re-entering the induction system 
and interfering with proper fuel metering operation.
VLA.040
    (a) Each exhaust system must ensure safe disposal of exhaust gases 
without fire hazard or carbon monoxide contamination in the personnel 
compartment.
    (b) Each exhaust system part with a surface hot enough to ignite 
flammable fluids or vapours must be located or shielded so that leakage 
from any system carrying flammable fluids or vapours will not result in 
a fire caused by impingement of the fluids or vapours on any part of 
the exhaust system including shields for the exhaust system.
    (c) Each exhaust system component must be separated by fireproof 
shields from adjacent flammable parts of the airplane that are outside 
the engine compartment.
    (d) No exhaust gases may discharge dangerously near any fuel or oil 
system drain.
    (e) Each exhaust system component must be ventilated to prevent 
points of excessively high temperature.
    (f) Each exhaust heat exchanger must incorporate means to prevent 
blockage of the exhaust port after any internal heat exchanger failure.
    (g) No exhaust gases may be discharged where they will cause a 
glare seriously affecting the pilot's vision at night.
VLA.045
    (a) The power or supercharger control must give a positive and 
immediate responsive means of controlling its engine or supercharger.
    (b) If a power control incorporates a fuel shut-off feature, the 
control must have a means to prevent the inadvertent movement of the 
control into the shut-off position. The means must--
    (1) Have a positive lock or stop at the idle position; and
    (2) Require a separate and distinct operation to place the control 
in the shut-off position.
    (c) Each power or thrust control must be designed so that if the 
control separates at the engine fuel metering device, the airplane is 
capable of continuing safe flight and landing.
VLA.050
    (a) The control must require a separate and distinct operation to 
move the control toward lean or shut-off position.
    (b) Each manual engine mixture control must be designed so that, if 
the control separates at the engine fuel metering device, the airplane 
is capable of continuing safe flight and landing.
VLA.055
    If warning, caution, or advisory lights are installed in the 
cockpit, they must be--
    (a) Red, for warning lights (lights indicating a hazard which may 
require immediate corrective action);
    (b) Amber, for caution lights (lights indicating the possible need 
for future corrective action);
    (c) Green, for safe operation lights; and
    (d) Any other color, including white, for lights not described in 
paragraphs (a) through (c) of this section, provided the

[[Page 70346]]

color differs sufficiently from the colors prescribed in paragraphs (a) 
through (c) of this section to avoid possible confusion.
    (e) If warning, caution, or advisory lights are installed in the 
cockpit, they must be effective under all probable cockpit lighting 
conditions.
VLA.060
    (a) Each instrument provided with static pressure case connections 
must be so vented that the influence of airplane speed, the opening and 
closing of windows, moisture, or other foreign matter, will not 
significantly affect the accuracy of the instruments.
    (b) The design and installation of a static pressure system must be 
such that--
    (1) Positive drainage of moisture is provided;
    (2) Chafing of the tubing, and excessive distortion or restriction 
at bends in the tubing, is avoided; and
    (3) The materials used are durable, suitable for the purpose 
intended, and protected against corrosion.
    (c) Each static pressure system must be calibrated in flight to 
determine the system error. The system error, in indicated pressure 
altitude, at sea-level, with a standard atmosphere, excluding 
instrument calibration error, may not exceed <plus-minus>9 m (<plus-
minus>30 ft) per 185 km/h (100 knots) speed for the appropriate 
configuration in the speed range between 1.3 V<INF>SO</INF> with flaps 
extended and 1.8 V<INF>S1</INF> with flaps retracted. However, the 
error need not be less than <plus-minus>9 m (<plus-minus>30 ft).
VLA.065
    For each airplane--
    (a) Each gyroscopic instrument must derive its energy from power 
sources adequate to maintain its required accuracy at any speed above 
the best rate-of-climb speed;
    (b) Each gyroscopic instrument must be installed so as to prevent 
malfunction due to rain, oil, and other detrimental elements; and
    (c) There must be a means to indicate the adequacy of the power 
being supplied to the instruments.
    (d) For Night VFR operation there must be at least two independent 
sources of power and a manual or an automatic means to select each 
power source for each instrument that uses a power source.
VLA.070
    (a) Electrical system capacity. Each electrical system must be 
adequate for the intended use. In addition--
    (1) Electric power sources, their transmission cables, and their 
associated control and protective devices, must be able to furnish the 
required power at the proper voltage to each load circuit essential for 
safe operation; and
    (2) Compliance with paragraph (a)(l) of this section must be shown 
by an electrical load analysis, or by electrical measurements, that 
account for the electrical loads applied to the electrical system in 
probable combinations and for probable durations.
    (b) Functions. For each electrical system, the following apply:
    (1) Each system, when installed, must be--
    (i) Free from hazards in itself, in its method of operation, and in 
its effects on other parts of the airplane;
    (ii) Protected from fuel, oil, water, other detrimental substances, 
and mechanical damage; and
    (iii) So designed that the risk of electrical shock to occupants 
and ground personnel is reduced to a minimum.
    (2) Electric power sources must function properly when connected in 
combination or independently.
    (3) No failure or malfunction of any electric power source may 
impair the ability of any remaining source to supply load circuits 
essential for safe operation.
    (4) Each electric power source control must allow the independent 
operation of each source, except that controls associated with 
alternators that depend on a battery for initial excitation or for 
stabilization need not break the connection between the alternator and 
its battery.
    (5) Each generator must have an overvoltage control designed and 
installed to prevent damage to the electrical system, or to equipment 
supplied by the electrical system, that could result if that generator 
were to develop an overvoltage condition.
    (d) Instruments. There must be a means to indicate to the pilot 
that the electrical power supplies are adequate for safe operation. For 
direct current systems, an ammeter in the battery feeder may be used.
    (e) Fire resistance. Electrical equipment must be so designed and 
installed that in the event of a fire in the engine compartment, during 
which the surface of the firewall adjacent to the fire is heated to 
1,100 [deg]C for 5 minutes or to a lesser temperature substantiated by 
the applicant, the equipment essential to continued safe operation and 
located behind the firewall will function satisfactorily and will not 
create an additional fire hazard. This may be shown by test or 
analysis.
    (f) External power. If provisions are made for connecting external 
power to the airplane, and that external power can be electrically 
connected to equipment other than that used for engine starting, means 
must be provided to ensure that no external power supply having a 
reverse polarity, or a reverse phase sequence, can supply power to the 
airplane's electrical system. The location must allow such provisions 
to be capable of being operated without hazard to the airplane or 
persons.
VLA.075
    (a) Each storage battery must be designed and installed as 
prescribed in this section.
    (b) Safe cell temperatures and pressures must be maintained during 
any probable charging and discharging condition. No uncontrolled 
increase in cell temperature may result when the battery is recharged 
(after previous complete discharge)--
    (1) At maximum regulated voltage or power;
    (2) During a flight of maximum duration; and
    (3) Under the most adverse cooling condition likely to occur in 
service.
    (c) Compliance with paragraph (b) of this section must be shown by 
tests unless experience with similar batteries and installations has 
shown that maintaining safe cell temperatures and pressures presents no 
problem.
    (d) No explosive or toxic gases emitted by any battery in normal 
operation, or as the result of any probable malfunction in the charging 
system or battery installation, may accumulate in hazardous quantities 
within the airplane.
    (e) No corrosive fluids or gases that may escape from the battery 
may damage surrounding structures or adjacent essential equipment.
    (f) Each nickel cadmium battery installation capable of being used 
to start an engine or auxiliary power unit must have provisions to 
prevent any hazardous effect on structure or essential systems that may 
be caused by the maximum amount of heat the battery can generate during 
a short circuit of the battery or of its individual cells.
    (g) Nickel cadmium battery installations capable of being used to 
start an engine or auxiliary power unit must have--
    (1) A system to control the charging rate of the battery 
automatically so as to prevent battery overheating;
    (2) A battery temperature sensing and over-temperature warning 
system with a means for disconnecting the battery

[[Page 70347]]

from its charging source in the event of an overtemperature condition; 
or
    (3) A battery failure sensing and warning system with a means for 
disconnecting the battery from its charging source in the event of 
battery failure.
    (h) In the event of a complete loss of the primary electrical power 
generating system, the battery must be capable of providing 30 minutes 
of electrical power to those loads that are essential to continued safe 
flight and landing. The 30-minute time period includes the time needed 
for the pilot(s) to recognize the loss of generated power and to take 
appropriate load shedding action.
VLA.080
    The instrument lights must--
    (a) Make each instrument and control easily readable and 
discernible;
    (b) Be installed so that their direct rays, and rays reflected from 
the windshield or other surface, are shielded from the pilot's eyes; 
and
    (c) Have enough distance or insulating material between current 
carrying parts and the housing so that vibration in flight will not 
cause shorting. (A cabin dome light is not an instrument light.)
VLA.085
    Each taxi and landing light must be designed and installed so 
that--
    (a) No dangerous glare is visible to the pilots;
    (b) The pilot is not seriously affected by halation;
    (c) It provides enough light for night operations; and
    (d) It does not cause a fire hazard in any configuration.
VLA.090
    (a) Electronic equipment and installations must be free from 
hazards in themselves, in their method of operation, and in their 
effects on other components.
    (b) For operations for which electronic equipment is required, 
compliance must be shown with CS-VLA 1309.
VLA.095
    (a) A placard meeting the requirements of this section must be 
installed on or near the magnetic direction indicator.
    (b) The placard must show the calibration of the instrument in 
level flight with the engine operating.
    (c) The placard must state whether the calibration was made with 
radio receivers on or off.
    (d) Each calibration reading must be in terms of magnetic headings 
in not more than 30[deg] increments.
    (e) If a magnetic non-stabilized direction indicator can have a 
deviation of more than 10[deg] caused by the operation of electrical 
equipment, the placard must state which electrical loads, or 
combination of loads, would cause a deviation of more than 10[deg] when 
turned on.
VLA.100
    The following placards must be plainly visible to the pilot:
    (a) A placard stating the following airspeeds (IAS):
    (1) Design maneuvering speed, V<INF>A</INF>;
    (2) The maximum landing gear operating speed, V<INF>LO</INF>.
    (b) A placard stating the following approved operation:
    (1) For day-VFR only operation, a placard stating, ``This airplane 
is classified as a very light airplane approved for day-VFR only, in 
non-icing conditions. All aerobatic maneuvers, including intentional 
spinning, are prohibited. See Flight Manual for other limitations.''
    (2) If night-VFR operation is approved, a placard stating, ``This 
airplane is classified as a very light airplane approved for day- and 
night-VFR operation, in non-icing conditions. All aerobatic maneuvers, 
including intentional spinning, are prohibited. See Flight Manual for 
other limitations.''
VLA.105
    (a) Airspeed limitations. The following information must be 
furnished--
    (1) Information necessary for the marking of the airspeed limits on 
the indicator, as required in CS-VLA 1545, and the significance of the 
color coding used on the indicator.
    (2) The speeds V<INF>A</INF>, V<INF>LO</INF>, V<INF>LE</INF> 
(maximum landing gear extended speed) where appropriate.
    (b) Weights. The following information must be furnished:
    (1) The maximum weight.
    (2) Any other weight limits, if necessary.
    (c) Center of gravity. The established c.g. limits required by CS-
VLA 23 must be furnished.
    (d) Maneuvers. Authorized maneuvers established in accordance with 
CS-VLA 3 must be furnished.
    (e) Flight load factors. Maneuvering load factors: the following 
must be furnished--
    (1) The factors corresponding to point A and point C in the figure 
for CS-VLA 333(b), stated to be applicable at V<INF>A</INF>.
    (2) The factors corresponding to point D and point E of figure 1 of 
CS-VLA 333(b) to be applicable at never exceed speed, V<INF>NE</INF>.
    (3) The factor with wing flaps extended as specified in CS-VLA 345.
    (f) The kinds of operation (day-VFR or day- and night-VFR, 
whichever is applicable) in which the airplane may be used, must be 
stated. The minimum equipment required for the operation must be 
listed.
    (g) Powerplant limitations. The following information must be 
furnished:
    (1) Limitation required by CS-VLA 1521.
    (2) Information necessary for marking the instruments required by 
CS-VLA 1549 through 1551.
    (3) Fuel and oil designation.
    (4) For two-stroke engines, fuel/oil ratio.
    (h) Placards. Placards required by CS-VLA 1555 through 1561 must be 
presented.

Increased Maximum Certificated Takeoff Weight and Increased Stall Speed

VLA.110
    If the maximum certificated takeoff weight is higher than 750 kg, 
but not more than 850 kg, the requirements in sections VLA.120 through 
VLA.210 apply.
VLA.115
    If the stall speed in landing configuration is higher than 45 
knots, but not more than 50 knots (CAS), the requirements in section 
VLA.120 through VLA.210 apply.
VLA.120
    The maximum horizontal distance traveled in still air, in km per 
1,000 m (nautical miles per 1,000 ft) of altitude lost in a glide, and 
the speed necessary to achieve this, must be determined with the engine 
inoperative and its propeller in the minimum drag position, and landing 
gear and wing flaps in the most favorable available position.
VLA.125
    (a) Each seat is to be equipped with at least a 4-point harness 
system;
    (b) The applicant shall evaluate the head strike path with 
validated methods, and minimize the risk of injury in case of a head 
contact with the aircraft structure or interior.
    (c) The design shall provide reasonable precautions to minimize the 
lumbar compression loads experienced by occupants in survivable crash 
landings;
    (d) Each seat/harness system shall be statically tested to an 
ultimate inertia load factor of 18g forward, considering an occupant's 
mass of 77 kg. The lapbelt should react 60% of this load, and the

[[Page 70348]]

upper torso restraint should react 40% of this load.
VLA.130
    (a) The airplane, although it may be damaged in emergency landing 
conditions, must be designed as prescribed in this section to protect 
each occupant under those conditions.
    (b) The structure must be designed to give each occupant reasonable 
chances of escaping injury in a minor crash landing when--
    (1) Proper use is made of seat belts and shoulder harnesses; and
    (2) The occupant experiences the ultimate inertia forces listed 
below:
    (i) Upward 3.0g
    (ii) Forward 9.0g
    (iii) Sideward 1.5g.
    (c) Each item of mass within the cabin that could injure an 
occupant if it came loose must be designed for the ultimate inertia 
load factors:
    (1) Upward, 3.0g;
    (2) Forward, 18.0g; and
    (3) Sideward, 4.5g.
    Engine mount and supporting structure are included in the above 
analysis if they are installed behind and above the seating 
compartment.
    (d) The structure must be designed to protect the occupants in a 
complete turnover, assuming, in the absence of a more rational 
analysis--
    (1) An upward ultimate inertia force of 3g; and
    (2) A coefficient of friction of 0.5 at the ground.
    (e) Each airplane with retractable landing gear must be designed to 
protect each occupant in a landing--
    (1) With the wheels retracted;
    (2) With moderate descent velocity; and
    (3) Assuming, in the absence of a more rational analysis;
    (i) A downward ultimate inertia force of 3g; and
    (ii) A coefficient of friction of 0.5 at the ground.
VLA.135
    (a) Each baggage compartment must be designed for its placarded 
maximum weight of contents and for the critical load distributions at 
the appropriate maximum load factors corresponding to the flight and 
ground load conditions for the airplane.
    (b) There must be means to prevent the contents of any baggage 
compartment from becoming a hazard by shifting, and to protect any 
controls, wiring, lines, equipment, or accessories whose damage of 
failure would affect safe operations.
    (c) Baggage compartments must be constructed of materials which are 
at least flame resistant.
    (d) Designs which provide for baggage to be carried must have means 
to protect the occupants from injury under the ultimate inertia forces 
specified in CS-VLA 561(b)(2).
    (e) If there is no structure between baggage and occupant 
compartments, the baggage items located behind the occupants and those 
which might become a hazard in a crash must be secured for 1.33 x 18g.
VLA.140
    (a) General. For each airplane, the following information must be 
furnished:
    (1) The takeoff distance determined under CS-VLA 51, the airspeed 
at the 15 m height, the airplane configuration (if pertinent), the kind 
of surface in the tests, and the pertinent information with respect to 
cowl flap position, use of flight path control devices, and use of the 
landing gear retraction system.
    (2) The landing distance determined under CS-VLA 75, the airplane 
configuration (if pertinent), the kind of surface used in the tests, 
and the pertinent information with respect to flap position and the use 
of flight path control devices.
    (3) The steady rate or gradient of climb determined under CS-VLA 65 
and 77, the airspeed, power, and the airplane configuration.
    (4) The calculated approximate effect on takeoff distance 
(paragraph (a)(1) of this section), landing distance (paragraph (a)(2) 
of this section), and steady rates of climb (paragraph (a)(3) of this 
section), of variations in altitude and temperature.
    (5) The maximum atmospheric temperature at which compliance with 
the cooling provisions of CS-VLA 1041 through 1047 is shown.
    (6) The glide performance determined under VLA.120.
    (b) Skiplanes. For skiplanes, a statement of the approximate 
reduction in climb performance may be used instead of new data for 
skiplane configuration, if--
    (1) The landing gear is fixed in both landplane and skiplane 
configurations;
    (2) The climb requirements are not critical; and
    (3) The climb reduction in the skiplane configurations is small 
(0.15 to 0.25 m/s (30 to 50 feet per minute)).
    (c) The following information concerning normal procedures must be 
furnished:
    (1) The demonstrated crosswind velocity and procedures and 
information pertinent to operation of the airplane in crosswinds, and
    (2) The airspeeds, procedures, and information pertinent to the use 
of the following airspeeds:
    (i) The recommended climb speed and any variation with altitude.
    (ii) V<INF>X</INF> (speed for best angle of climb) and any 
variation with altitude.
    (iii) The approach speeds, including speeds for transition to the 
balked landing condition.
    (d) An indication of the effect on takeoff distance of a grass 
surface as determined from at least one takeoff measurement on short 
mown dry grass must be furnished.
VLA.145
    (a) The rotation speed V<INF>R</INF>, is the speed at which the 
pilot makes a control input with the intention of lifting the airplane 
out of contact with the runway.
    (b) V<INF>R</INF> must not be less than stalling speed, 
V<INF>S1</INF>.
    (c) The Airplane Flight Manual must provide the rotation speed 
established above for normal takeoff procedures.
    If an Equivalent Level of Safety (ELOS) to CS-VLA 1143(g) and CS-
VLA 1147(b) is requested, VLA.150 and VLA.155 are applicable.
VLA.150
    Power or supercharger control attachment design must include:
    (a) Features which are not likely to separate in flight (i.e., a 
large load-bearing washer adjacent to the outside face of the power 
control cable rod end fitting which attaches to the fuel-metering 
device);
    (b) Mandatory inspection intervals;
    (c) Inspection procedures;
    (d) Component replacement criteria.
VLA.155
    Mixture control attachment design must include:
    (a) Features which are not likely to separate in flight (i.e., a 
large load-bearing washer adjacent to the outside face of the power 
control cable rod end fitting which attaches to the fuel-metering 
device);
    (b) Mandatory inspection intervals;
    (c) Inspection procedures;
    (d) Component replacement criteria.
VLA.160
    (a) For an airplane with independently controlled roll and 
directional controls, it must be possible to produce and to correct 
roll by unreversed use of the rolling control and to produce and to 
correct yaw by unreversed use of the directional control, up to the 
time the airplane stalls.
    (b) For an airplane with interconnected lateral and directional 
controls (2 controls) and for an airplane with only one of these 
controls, it must be possible to produce and correct roll

[[Page 70349]]

by unreversed use of the rolling control without producing excessive 
yaw, up to the time the airplane stalls.
    (c) The wing level stall characteristics of the airplane must be 
demonstrated in flight as follows: The airplane speed must be reduced 
with the elevator control until the speed is slightly above the 
stalling speed, then the elevator control must be pulled back so that 
the rate of speed reduction will not exceed 1.9 km/h (one knot) per 
second until a stall is produced, as shown by an uncontrollable 
downward pitching motion of the airplane, or until the control reaches 
the stop. Normal use of the elevator control for recovery is allowed 
after the control has been held against the stop for not less than two 
seconds.
    (d) Except where made inapplicable by the special features of a 
particular type of airplane, the following apply to the measurement of 
loss of altitude during a stall:
    (1) The loss of altitude encountered in the stall (power on or 
power off) is the change in altitude (as observed on the sensitive 
altimeter testing installation) between the altitude at which the 
airplane pitches and the altitude at which horizontal flight is 
regained.
    (2) If power or thrust is required during stall recovery, the power 
or thrust used must be that which would be used under the normal 
operating procedures selected by the applicant for this maneuver. 
However, the power used to regain level flight may not be applied until 
flying control is regained.
    (e) During the recovery part of the maneuver, it must be possible 
to prevent more than 15[deg] of roll or yaw by the normal use of 
controls.
    (f) Compliance with the requirements of this section must be shown 
under the following conditions:
    (1) Wing flaps. Retracted, fully extended and each intermediate 
normal operating position;
    (2) Landing gear. Retracted and extended;
    (3) Cowl flaps. Appropriate to configuration;
    (4) Power
    (i) Power off; and
    (ii) 75% maximum continuous power. If the power-to-weight ratio at 
75% of maximum continuous power results in extreme nose-up attitudes, 
the test may be carried out with the power required for level flight in 
the landing configuration at maximum landing weight and a speed of 1.4 
stalling speed, V<INF>S0</INF>, but the power may not be less than 50% 
maximum continuous power.
    (5) Trim. The airplane trimmed at a speed as near 1.5 
V<INF>S1</INF> as practicable.
    (6) Propeller. Full increase rpm position for the power off 
condition.
VLA.165
    Turning flight and accelerated stalls must be demonstrated in tests 
as follows:
    (a) Establish and maintain a coordinated turn in a 30[deg] bank. 
Reduce speed by steadily and progressively tightening the turn with the 
elevator until the airplane is stalled or until the elevator has 
reached its stop. The rate of speed reduction must be constant, and--
    (1) For a turning flight stall, may not exceed 1.9 km/h (one knot) 
per second; and
    (2) For an accelerated stall, be 5.6 to 9.3 km/h (3 to 5 knots) per 
second with steadily increasing normal acceleration.
    (b) When the stall has fully developed or the elevator has reached 
its stop, it must be possible to regain level flight by normal use of 
controls and without--
    (1) Excessive loss of altitude;
    (2) Undue pitchup;
    (3) Uncontrollable tendency to spin;
    (4) Exceeding 60[deg] of roll in either direction from the 
established 30[deg] bank; and
    (5) For accelerated entry stalls, without exceeding the maximum 
permissible speed or the allowable limit load factor.
    (c) Compliance with the requirements of this section must be shown 
with--
    (1) Wing Flaps. Retracted and fully extended for turning flight and 
accelerated entry stalls, and intermediate, if appropriate, for 
accelerated entry stalls;
    (2) Landing Gear. Retracted and extended;
    (3) Cowl Flaps. Appropriate to configuration;
    (4) Power. 75% maximum continuous power. If the power-to-weight 
ratio at 75% of maximum continuous power results in extreme nose-up 
attitudes, the test may be carried out with the power required for 
level flight in the landing configuration at maximum landing weight and 
a speed of 1.4 V<INF>S0</INF>, but the power may not be less than 50% 
maximum continuous power.
    (5) Trim. 1.5 V<INF>S1</INF> or minimum trim speed, whichever is 
higher.
VLA.170
    (a) Three-control airplanes. The stability requirements for three-
control airplanes are as follows:
    (1) The static directional stability, as shown by the tendency to 
recover from a skid with the rudder free, must be positive for any 
landing gear and flap position appropriate to the takeoff, climb, 
cruise, and approach configurations. This must be shown with power up 
to maximum continuous power, and at speeds from 1.2 V<INF>S1</INF> up 
to maximum allowable speed for the condition being investigated. The 
angle of skid for these tests must be appropriate to the type of 
airplane. At larger angles of skid up to that at which full rudder is 
used or a control force limit in CS-VLA 143 is reached, whichever 
occurs first, and at speeds from 1.2 V<INF>S1</INF> to V<INF>A</INF>, 
the rudder pedal force must not reverse.
    (2) The static lateral stability, as shown by the tendency to raise 
the low wing in a slip, must not be negative for any landing gear and 
flap positions. This must be shown with power up to 75% of maximum 
continuous power at speeds above 1.2 V<INF>S1</INF>, up to the maximum 
allowable speed for the configuration being investigated. The static 
lateral stability may not be negative at 1.2 V<INF>S1</INF>. The angle 
of slip for these tests must be appropriate to the type of airplane, 
but in no case may the slip angle be less than that obtainable with 
10[deg] of bank.
    (3) In straight, steady slips at 1.2 V<INF>S1</INF> for any landing 
gear and flap positions, and for power conditions up to 50% of maximum 
continuous power, the rudder control movements and forces must increase 
steadily (but not necessarily linearly) as the angle of slip is 
increased up to the maximum appropriate to the type of airplane. At 
larger slip angles up to the angle at which full rudder or aileron 
control is used or a control force limit contained in CS-VLA 143 is 
obtained, aileron control movements and forces must not reverse. Enough 
bank must accompany slipping to hold a constant heading. Rapid entry 
into, or recovery from, a maximum slip may not result in uncontrollable 
flight characteristics. The applicant must demonstrate that lateral 
static stability characteristics do not result in any unsafe handling 
qualities.
    (b) Two-control (or simplified control) airplanes. The stability 
requirements for two-control airplanes are as follows:
    (1) The directional stability of the airplane must be shown by 
showing that, in each configuration, it can be rapidly rolled from a 
45[deg] bank in one direction to a 45[deg] bank in the opposite 
direction without showing dangerous skid characteristics.
    (2) The lateral stability of the airplane must be shown by showing 
that it will not assume a dangerous attitude or speed when the controls 
are abandoned for 2 minutes. This must be done in moderately smooth air 
with the airplane trimmed for straight level flight at 0.9 
V<INF>H</INF> (maximum speed in level flight with maximum continuous 
power) or V<INF>C</INF>

[[Page 70350]]

(design cruising speed), whichever is lower, with flaps and landing 
gear retracted, and with a rearward center of gravity.
    If an ELOS to CS-VLA 161(b)(2)(ii) is requested, VLA.175 through 
VLA.210 are applicable.
VLA.175
    Longitudinal trim. The airplane must maintain longitudinal trim 
under each of the following conditions:
    (a) Approach with landing gear extended and with--
    (i) A 3[deg] angle of descent, with flaps retracted and at a speed 
of 1.4 V<INF>S1</INF>;
    (ii) A 3[deg] angle of descent, flaps in the landing position(s) at 
reference landing approach speed, V<INF>REF</INF>; and
    (iii) An approach gradient equal to the steepest used in the 
landing distance demonstrations of CS 23.75, flaps in the landing 
position(s) at V<INF>REF</INF>.
VLA.180
    For normal, utility and aerobatic category reciprocating engine-
powered airplanes of 2,722 kg (6,000 lb) or less maximum weight, the 
reference landing approach speed, V<INF>REF</INF>, must not be less 
than the greater of minimum control speed, V<INF>MC</INF>, determined 
under CS 23.149(b) with the wing flaps in the most extended takeoff 
setting, and 1.3 V<INF>SO</INF>.
VLA.185
    (a) A steady approach at not less than V<INF>REF</INF>, determined 
in accordance with CS 23.73(a), (b) or (c) as appropriate, must be 
maintained down to 15 m (50 ft) height and--
    (1) The steady approach must be at a gradient of descent not 
greater than 5.2% (3[deg]) down to the 15 m (50 ft) height.
    (b) A constant configuration must be maintained throughout the 
maneuver.
    (c) The landing must be made without excessive vertical 
acceleration or tendency to bounce, nose-over, ground loop, porpoise, 
or water loop.
    (d) It must be shown that a safe transition to the balked landing 
conditions of CS 23.77 can be made from the conditions that exist at 
the 15 m (50 ft) height, at maximum landing weight, or the maximum 
landing weight for altitude and temperature of CS 23.63(c)(2) or 
(d)(2), as appropriate.
VLA.190
    (a) Each normal, utility, and aerobatic category reciprocating 
engine-powered airplane of 2,722 kg (6,000 lb) or less maximum weight 
must be able to maintain a steady gradient of climb at sea-level of at 
least 3.3% with--
    (1) Takeoff power on each engine;
    (2) The landing gear extended;
    (3) The wing flaps in the landing position, except that if the 
flaps may safely be retracted in 2 seconds or less without loss of 
altitude and without sudden changes of angle of attack, they may be 
retracted; and
    (4) A climb speed equal to V<INF>REF</INF>, as defined in CS 
23.73(a).
VLA.195
    (a) It must be possible to carry out the following maneuvers 
without requiring the application of single-handed control forces 
exceeding those specified in CS 23.143(c), unless otherwise stated. The 
trimming controls must not be adjusted during the maneuvers:
    (1) With power off, landing gear and flaps extended and the 
airplane as nearly as possible in trim at V<INF>REF</INF>, obtain and 
maintain airspeeds between 1.1 V<INF>S0</INF> and either 1.7 
V<INF>S0</INF> or V<INF>FE</INF> (maximum flap extended speed), 
whichever is lower, without requiring the application of two-handed 
control forces exceeding those specified in CS 23.143(c).
    (b) It must be possible, with a pilot control force of not more 
than 44.5 N (10 lbf), to maintain a speed of not more than 
V<INF>REF</INF> during a power-off glide with landing gear and wing 
flaps extended.
VLA.200
    It must be possible, while in the landing configuration, to safely 
complete a landing without exceeding the one-hand control force limits 
specified in CS 23.143(c) following an approach to land--
    (a) At a speed of V<INF>REF</INF> 9.3 km/h (5 knots);
    (b) With the airplane in trim, or as nearly as possible in trim and 
without the trimming control being moved throughout the maneuver;
    (c) At an approach gradient equal to the steepest used in the 
landing distance demonstration of CS 23.75;
    (d) With only those power changes, if any, which would be made when 
landing normally from an approach at V<INF>REF</INF>.
VLA.205
    (a) Approach--It must be possible using a favorable combination of 
controls, to roll the airplane from a steady 30[deg] banked turn 
through an angle of 60[deg], so as to reverse the direction of the turn 
within--
    (1) For an airplane of 2,722 kg (6,000 lb) or less maximum weight, 
4 seconds from initiation of roll; and
    (2) For an airplane of over 2,722 kg (6,000 lb) maximum weight, 
1,000/W + 1,300 but not more than 7 seconds, where W is weight in kg. 
(W + 2800/2200 but not more than 7 seconds where W is weight in lb.).
    (b) The requirement of paragraph (a) of this section must be met 
when rolling the airplane in each direction in the following 
conditions--
    (1) Flaps in the landing position(s);
    (2) Landing gear extended;
    (3) All engines operating at the power for a 3[deg] approach; and
    (4) The airplane trimmed at V<INF>REF</INF>.
VLA.210
    (a) Landing. The stick force curve must have a stable slope at 
speeds between 1.1 V<INF>S1</INF> and 1.8 V<INF>S1</INF> with--
    (1) Flaps in the landing position;
    (2) Landing gear extended; and
    (3) The airplane trimmed at--
    (i) V<INF>REF</INF>, or the minimum trim speed if higher, with 
power off; and
    (ii) V<INF>REF</INF> with enough power to maintain a 3[deg] angle 
of descent.

Rechargeable Lithium Ion Battery

VLA.215
    The applicant must consider the following safety objectives when 
showing compliance with regulations applicable to the rechargeable 
lithium ion battery.
    Each rechargeable lithium ion battery installation must:
    (a) Be designed to maintain safe cell temperatures and pressures 
under all foreseeable operating conditions to prevent fire and 
explosion;
    (b) Be designed to prevent the occurrence of self-sustaining, 
uncontrollable increases in temperature or pressure, and automatically 
control the charge rate of each cell to protect against adverse 
operating conditions, such as cell imbalance, back charging, 
overcharging, and overheating;
    (c) Not emit explosive or toxic gases, either in normal operation 
or as a result of its failure, that may accumulate in hazardous 
quantities within the airplane;
    (d) Meet the requirements of 14 CFR 23.2325(g);
    (e) Not damage surrounding structure or adjacent systems, 
equipment, components, or electrical wiring from corrosive or any other 
fluids or gases that may escape in such a way as to cause a major or 
more-severe failure condition;
    (f) Have provisions to prevent any hazardous effect on airplane 
structure or systems caused by the maximum amount of heat it can 
generate due to any failure of it or its individual cells;
    (g) Have a failure sensing and warning system to alert the 
flightcrew if its failure affects safe operation of the airplane;
    (h) Have a monitoring and warning feature that alerts the 
flightcrew when

[[Page 70351]]

its charge state falls below acceptable levels if its function is 
required for safe operation of the airplane;
    (i) Have a means to disconnect from its charging source in the 
event of an over-temperature condition, cell failure, or battery 
failure.

    Issued in Kansas City, Missouri, on October 5, 2023.
Patrick R. Mullen,
Manager, Technical Policy Branch, Policy and Standards Division, 
Aircraft Certification Service.
[FR Doc. 2023-22492 Filed 10-10-23; 8:45 am]
BILLING CODE 4910-13-P


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