CIVIL AIR REGULATIONS CIVIL AERONAUTICS BOARD … More Navion Files/car_part3.pdf · Stalls 3.120...

CIVIL AIR REGULATIONS——————

PART 3——AIRPLANEAIRWORTHINESS——NORMAL,UTILITY, ACROBATIC, AND

RESTRICTED PURPOSECATEGORIES

——————As amended to November 1, 1949

CIVIL AERONAUTICS BOARD

WASHINGTON, D.C.

S UBPART A—A IRWORTHINESS REQUIREMENTS

GeneralSec.3.1 Scope3.2 Date of effectiveness.

AIRPLANE CATEGORIES

3.6 Airplane categories.

AIRWORTHINESS CERTIFICATES

3.11 Classification.

Type Certificates3.15 Requirements for issuance.3.16 Data required for NC and NR certification.3.17 Inspection and tests for NC and NR

certification.3.18 Inspection.3.19 Flight tests.

Changes3.23 Changes.3.24 Minor changes.3.25 Major changes.3.26 Service experience changes.3.27 Application to earlier airworthiness

requirements.

Approval of Materials, Parts, Processes, andAppliances

3.31 Specifications.

Definitions3.41 Standard atmosphere.3.42 Hot-day condition.3.43 Airplane configuration.3.44 Weights.3.45 Power.3.46 Speeds.3.47 Structural terms.3.48 Susceptibility of materials to fire.

Subpart B—Flight RequirementsGeneral

3.61 Policy re proof of compliance.3.62 Flight test pilot.

3.61 Policy re proof of compliance.3.63 Noncompliance with test requirements.3.64 Emergency egress.3.65 Report.

Weight Range and Center of Gravity3.71 Weight and balance.3.72 Use of ballast.3.73 Empty weight.3.74 Maximum weight.3.75 Minimum weight.3.76 Center of gravity position.

Performance Requirements General

3.81 Performance.3.82 Definition of stalling speeds.3.83 Stalling speed.

Take-off3.84 Take-off.

Climb3.85 Climb.

Landing3.86 Landing.

Flight Characteristics3.105 Requirements.

Controllability3.106 General.3.107-U Approved acrobatic maneuvers.3.108-A Acrobatic maneuvers.3.109 Longitudinal control.3.110 Lateral and directional control.3.111 Minimum control speed (Vmc ).

Trim3.112 Requirements.

Stability3.113 General.3.114 Static longitudinal stability.3.115 Specific conditions.3.116 Instrument stick force measurements.3.117 Dynamic longitudinal stability.3.118 Directional and lateral stability.

Stalls3.120 Stalling demonstration.3.121 Climbing stalls.3.122 Turning flight stalls.3.123 One-engine-inoperative stalls.

Spinning3.124 Spinning.

Ground and Water Characteristics3.143 Requirements.3.144 Longitudinal stability and control.3.145 Directional stability and control.3.146 Shock absorption.3.147 Spray characteristics.

Flutter and Vibration3.159 Flutter and vibration.

Subpart C—Strength RequirementsGeneral

3.171 Loads.3.172 Factor of safety.3.173 Strength and deformations.3.174 Proof of structure.

Flight Loads3.181 General.3.182 Definition of flight load factor.

Symmetrical Flight Conditions(Flaps Retracted)

3.183 General.3.184 Design air speeds.3.185 Maneuvering envelope.3.186 Maneuvering load factors.3.187 Gust envelope.3.188 Gust load factors.3.189 Airplane equilibrium.

Flaps Extended Flight Conditions3.190 Flaps extended flight conditions.

Unsymmetrical Flight Conditions3.191 Unsymmetrical flight conditions.

Supplementary Conditions3.194 Special condition for rear lift truss.3.195 Engine torque effects.3.196 Side load on engine mount.

Control Surface Loads3.211 General.3.212 Pilot effort.3.213 Trim tab effects.

Horizontal Tail Surfaces3.214 Horizontal tail surfaces.3.215 Balancing loads.3.216 Maneuvering loads.

3.214 Horizontal tail surfaces.3.217 Gust loads.3.218 Unsymmetrical loads.

Vertical Tail Surface3.219 Maneuvering loads.3.220 Gust loads.3.221 Outboard fins.

Ailerons, Wing Flaps, Tabs, Etc.3.222 Ailerons.3.223 Wing flaps.3.224 Tabs.3.225 Special devices.

Control System Loads3.231 Primary flight controls and systems.3.232 Dual controls.3.233 Ground gust conditions.3.234 Secondary controls and systems.

Ground Loads3.241 Ground loads.3.242 Design weight.3.243 Load factor for landing conditions.

Landing Cases and Attitudes3.244 Landing cases and attitudes.3.245 Level landing.3.246 Tail down.3.247 One-wheel landing.

Ground Roll Conditions3.248 Braked roll.3.249 Side load.

Tail Wheels3.250 Supplementary conditions for tail wheels.3.251 Obstruction load.3.252 Side load.

Nose Wheels3.253 Supplementary conditions for nose

wheels.3.254 Aft load.3.255 Forward load.3.256 Side load.

Skiplanes3.257 Supplementary conditions for skiplanes.

Water Loads3.265 General.

Design Weight3.266 Design weight.

Boat Seaplanes3.267 Local bottom pressures.3.268 Distributed bottom pressures.

3.267 Local bottom pressures.3.269 Step loading condition.3.270 Bow loading condition.3.271 Stern loading condition.3.272 Side loading condition.

Float Seaplanes3.273 Landing with inclined reactions.3.275 Landing with vertical reactions.3.277 Landing with side load.3.278 Supplementary load conditions.3.279 Bottom loads.

Wing-Tip Float and Sea Wing Loads3.280 Wing-tip float loads.3.281 Wing structure.3.282 Sea wing loads.

Subpart D—Design and ConstructionGeneral

3.291 General.3.292 Materials and workmanship.3.293 Fabrication methods.3.294 Standard fastenings.3.295 Protection.3.296 Inspection provisions.

Structural Parts3.301 Material strength properties and

design values.3.302 Special factors.3.303 Variability factor.3.304 Castings.3.305 Bearing factors.3.306 Fitting factor.3.307 Fatigue strength.

Flutter and Vibration3.311 Flutter and vibration prevention

measures.

Wings3.317 Proof of strength.3.318 Ribs.3.319 External bracing.3.320 Covering.

Control Surfaces (Fixed and Movable)3.327 Proof of strength.3.328 Installation.3.329 Hinges.

Control Systems3.335 General.3.336 Primary flight controls.3.337 Trimming controls.3.338 Wing flap controls.3.339 Flap interconnection.3.340 Stops.3.341 Control system locks.3.342 Proof of strength.3.343 Operation test.

Control System Details3.344 General.

3.345 Cable systems.3.346 Joints.3.347 Spring devices.

Landing GearShock Absorbers

3.351 Tests.3.352 Shock absorption tests.3.353 Limit drop tests.3.354 Limit load factor determination.3.355 Reserve energy absorption drop tests.

Retracting Mechanism3.356 General.3.357 Emergency operation.3.358 Operation test.3.359 Position indicator and warning device.3.360 Control.

Wheels and Tires3.361 Wheels.3.362 Tires.

Brakes3.363 Brakes.

Skis3.364 Skis.3.365 Installation3.366 Tests.

Hulls and Floats3.371 Buoyancy (main seaplane floats).3.372 Buoyancy (boat seaplanes).3.373 Water stability.

FuselagePilot Compartment

3.381 General.3.382 Vision.3.383 Pilot windshield and windows.3.384 Cockpit controls.3.385 Instruments and markings.

Emergency Provisions3.386 Protection.3.387 Exits.3.388 Fire precautions.

Personnel and Cargo Accommodations3.389 Doors.3.390 Seats and berths.3.391 Safety belt or harness provisions.3.392 Cargo compartments.3.393 Ventilation.

Miscellaneous3.401 Leveling marks.

SUBPART E ——POWER-PLANTINSTALLATIONS; RECIPROCATING

ENGINESGeneral

3.411 Components.

Engines and Propellers3.415 Engines.3.416 Propellers.3.417 Propeller vibration.3.418 Propeller pitch and speed limitations.3.419 Speed limitations for fixed pitch

propellers, ground adjustable pitchpropellers, and automatically varyingpitch propellers which cannot becontrolled in flight.

3.420 Speed and pitch limitations forcontrollable pitch propellers withoutconstant speed controls.

3.421 Variable pitch propellers with constantspeed controls.

3.422 Propeller clearance.

Fuel System3.429 General.

Arrangement3.430 Fuel system arrangement.3.431 Multiengine fuel system arrangement.3.432 Pressure cross feed arrangements.

Operation3.433 Fuel flow rate.3.434 Fuel flow rate for gravity feed systems.3.435 Fuel flow rate for pump systems.3.436 Fuel flow rate for auxiliary fuel systems

and fuel transfer systems.3.437 Determination of unusable fuel supply

and fuel system operation on low fuel.3.438 Fuel system hot weather operation.3.439 Flow between interconnected tanks.

Fuel Tanks3.440 General.3.441 Fuel tank tests.3.442 Fuel tank installation.3.443 Fuel tank expansion space.3.444 Fuel tank sump.3.445 Fuel tank filler connection.3.446 Fuel tank vents and carburetor vapor

vents.3.447-A Fuel tank vents.3.448 Fuel tank outlet.

Fuel Pumps3.449 Fuel pump and pump installation.

Lines, Fittings, and Accessories3.550 Fuel system lines, fittings, and

accessories.3.551 Fuel valves.3.552 Fuel strainer.

Drains and Instruments3.553 Fuel system drains.3.554 Fuel system instruments.

Oil System3.561 Oil system.3.562 Oil cooling.

Oil Tanks3.563 Oil tanks.3.564 Oil tank tests.3.565 Oil tank installation.3.566 Oil tank expansion space.3.567 Oil tank filler connection.3.568 Oil tank vent.3.569 Oil tank outlet.

Lines, Fittings, and Accessories3.570 Oil system lines, fittings, and accessories.3.571 Oil valves.3.572 Oil radiators.3.573 Oil filters.3.574 Oil system drains.3.575 Engine breather lines.3.576 Oil system instruments.3.577 Propeller feathering system.

Cooling3.581 General.

Tests3.582 Cooling tests.3.583 Maximum anticipated summer air

temperatures.3.584 Correction factor for cylinder head, oil

inlet, carburetor, air, and engine coolantinlet temperatures.

3.585 Correction factor for cylinder barreltemperatures.

3.586 Cooling test procedure for single-engineairplanes.

3.587 Cooling test procedure for multiengineairplanes.

Liquid Cooling Systems3.588 Independent systems.3.589 Coolant tank.3.590 Coolant tank tests.3.591 Coolant tank installation.3.592 Coolant tank filler connection.3.593 Coolant lines, fittings, and

accessories.

3.594 Coolant radiators.3.595 Cooling system drains.3.596 Cooling system instruments.

Induction System3.605 General.3.606 Induction system de-icing and anti-

icing provisions.3.607 Carburetor de-icing fluid flow rate.3.608 Carburetor fluid de-icing system

capacity.3.609 Carburetor fluid de-icing system detail

design.3.610 Carburetor air preheater design.3.611 Induction system ducts.3.612 Induction system screens.

Exhaust System3.615 General.3.616 Exhaust manifold.3.617 Exhaust heat exchangers.3.618 Exhaust heat exchangers used in

ventilating air heating systems.

Fire Wall and Cowling3.623 Fire walls.3.624 Fire wall construction.3.625 Cowling.

Power-Plant Controls and Accessories Controls3.627 Power-plant controls.3.628 Throttle controls.3.629 Ignition switches.3.630 Mixture controls.3.631 Propeller speed and pitch controls.3.632 Propeller feathering controls.3.633 Fuse system controls.3.634 Carburetor air preheat controls.

Accessories3.635 Power-plant accessories.3.636 Engine battery ignition systems.

Power-Plant Fire Protection3.637 Power-plant fire protection.

Subpart F- Equipment3.651 General.3.652 Functional and installational

requirements.

Basic Equipment3.655 Required basic equipment.

Instruments; Installation

General3.661 Arrangement and visibility of instrument

installations.3.662 Instrument panel vibration characteristics.

Flight and Navigational Instruments3.663 Air-speed indicating system.3.664 Air-speed indicator marking.3.665 Static air vent system.3.666 Magnetic direction indicator.3.667 Automatic pilot system.3.668 Gyroscopic indicators (air-driven type).3.669 Suction gauge.

Power-Plant Instruments3.670 Operational markings.3.671 Instrument lines.3.672 Fuel quantity indicator.3.673 Fuel flowmeter system.3.674 Oil quantity indicator.3.675 Cylinder head temperature indicating

system for air-cooled engines.3.676 Carburetor air temperature indicating

system.

Electrical Systems and Equipment3.681 Installation.

Batteries3.682 Batteries.3.683 Protection against acid.3.684 Battery vents.

Generators3.685 Generator.3.686 Generator controls.3.687 Reverse current cut-out.

Master Switch3.688 Arrangement.3.689 Master switch installation.

Protective Devices3.690 Fuses or circuit breakers.3.691 Protective devices installation.3.692 Spare fuses.

Electric Cables3.693 Electric cables.

Switches3.694 Switches.3.695 Switch installation.

Instrument Lights3.696 Instrument lights.3.697 Instrument light installation.

Landing Lights3.698 Landing lights.3.699 Landing light installation.

Position Lights3.700 Type.3.701 Forward position light installation.3.702 Rear position light installation.

3.703 Flashing rear position lights.

Anchor Lights3.704 Anchor light.3.705 Anchor light installation.

Safety Equipment; Installation3.711 Marking.3.712 De-icers.3.713 Flare requirements.3.714 Flare installation.3.715 Safety belts.

Emergency Flotation and Signaling Equipment3.716 Rafts and life preservers.3.717 Installation.3.718 Signaling device.

Radio Equipment; Installation3.721 General.

Miscellaneous Equipment; Installation3.725 Accessories for multiengine airplanes.

Hydraulic Systems3.726 General.3.727 Tests.3.728 Accumulators.

Subpart G - Operating Limitation andInformation

3.735 General.

Limitations3.737 Limitations.

Air Speed3.738 Air speed.3.739 Never exceed speed (Vne).3.740 Maximum structural cruising speed

(Vno).3.741 Maneuvering speed (Vp).3.742 Flaps-extended speed (Vfe).3.743 Minimum control speed (Vmc).

Power Plant3.744 Power plant.3.745 Take-off operation.3.746 Maximum continuous operation.3.747 Fuel octane rating.

Airplane Weight3.748 Airplane weight.

Minimum Flight Crew3.749 Minimum flight crew.

Types of Operation3.750 Types of operation.

Markings and Placards3.755 Markings and placards.

Instrument Markings3.756 Instrument markings.

3.757 Air-speed indicator.3.758 Magnetic direction indicator.3.759 Power-plant instruments.3.760 Oil quantity indicators.3.761 Fuel quantity indicator.

Control Markings3.762 General.3.763 Aerodynamic controls.3.764 Power-plant fuel controls.3.765 Accessory and auxiliary controls.

Miscellaneous3.766 Baggage compartments, ballast location,

and special seat loading limitations.3.767 Fuel, oil, and coolant filler openings.3.768 Emergency exit placards.3.769 Approved flight maneuvers.3.770 Airplane category placard.

Airplane Flight Manual3.777 Airplane Flight Manual.3.778 Operating limitations.3.779 Operating procedures.3.780 Performance information.

Subpart H - Identification Data3.791 Name plate.3.792 Airworthiness certificate number.

AUTHORITY: §§ 3.1 to 3.792 issued undersec. 205(a), 52 Stat. 984; 49 U. S. C. 425(a).Interpret or apply secs. 601; 52 Stat. 1007; 49U.S.C. 551.

SOURCE: §§ 3.1 to 3.792 contained inAmendment 03-0, Civil Air Regulations, 11 F.R.13368, except as noted following sectionsaffected. Redesignated at 13 F.R. 5486.

SUBPART A—AIRWORTHINESSREQUIREMENTS

GENERAL

§ 3.1 Scope. An airplane which has nofeatures or characteristics rendering it unsafe forthe category for which it is to be certificated iseligible for type and airworthiness certification, ifit complies with all applicable provisions of thispart, or, in the event it does not so comply, if it isshown to meet the same level of safety as thatprovided for in this part.

§ 3.2 Date of effectiveness. (a) Airplanescertificated as a type on or after November 13,1945, shall comply either with (1) the entireprovisions of Part 4a of this chapter in effectimmediately prior to November 9, 1945, or (2) theentire provisions prescribed in this part, exceptthat airplanes certificated under (1) may

incorporate provisions of (2) when theAdministrator finds the standard of safety to beequivalent to the particular and all related itemsof the latter.

(b) Airplanes certificated as a type on orafter January 1, 1947, shall comply with theprovisions contained in this part. If theprototype is not flown prior to January 1, 1947,and satisfactory evidence is presented indicatingthat the design work of the type was welladvanced prior to November 13, 1945, and thedelay of completion of the airplane was due tocauses beyond the manufacturer’s control, theAdministrator may certificate the airplane as atype under the provisions of Part 4a of thischapter which were in effect prior to November 9,1945.

(c) Unless otherwise specified, compliancewith an amendment to this part shall bemandatory only for airplanes for whichapplication for a type certificate has beenreceived subsequent to the effective date of suchamendment.

AIRPLANE CATEGORIES

§ 3.6 Airplane categories. (a) In this partairplanes are divided upon the basis of theirintended operation into the following categoriesfor the purpose of certification.

(1) Normal—Suffix"N". Airplanes in thiscategory are intended for nonacrobatic,nonscheduled passenger, and nonscheduledcargo operation.

(2) Utility—Suffix "U". Airplanes in thiscategory are intended for normal operations andlimited acrobatic maneuvers. These airplanes arenot suited for use in snap or inverted maneuvers.

NOTE: The following interpretation ofparagraph (a) (2) was issued May 15, 1947, 12F.R. 3434:

The phrase "limited acrobatic maneuvers" asused in § 3.6 is interpreted to include steep turns,spins, stalls (except whip stalls), lazy eights, andchandelles.

(3) Acrobatic—Suffix "A". Airplanes inthis category will have no specific restrictions asto type of maneuver permitted unless thenecessity therefor is disclosed by the requiredflight tests.

(4) Restricted purpose—Suffix "R".Airplanes in this category are intended to beoperated for restricted purposes not logicallyencompassed by the foregoing categories. Therequirements of this category shall consist of allof the provisions for any one of the foregoingcategories which are not rendered inapplicable bythe nature of the special purpose involved, plussuitable operating restrictions which theAdministrator finds will provide a level of safetyequivalent to that contemplated for the foregoingcategories.

(b) An airplane may be certificated underthe requirements of a particular category, or inmore than one category, provided that all of therequirements of such categories are met.Sections of this part which apply to only one ormore, but not all, categories are identified in thispart by the appropriate suffixes, as indicatedabove, added to the section number. All sectionsnot identified by a suffix are applicable to allcategories except as otherwise specified.

NOTE: For rules governing the eligibility ofairplanes certificated under this part for use in aircarrier operations see Parts 40, 41, 42, and 61 ofthis chapter.

AIRWORTHINESS CERTIFICATES

§ 3.11 Classification. (a) Airworthinesscertificates are classified as follows:

(1) NC (standard) certificates. In order tobecome eligible for an NC (standard) certificate,an airplane must be shown to comply with therequirements contained in this part for at leastone category, but not the restricted-purposecategory.

(2) NR (restricted) certificates. In order tobecome eligible for an NR (restricted) certificate,an airplane must be shown to comply with therequirements of the restricted purpose category.

(3) NX (experimental) certificates. Anairplane will become eligible for an NX(experimental) certificate when the applicantpresents satisfactory evidence that the airplane isto be flown for experimental purposes and theAdministrator finds it may, with appropriaterestrictions, be operated for that purpose in amanner which does not endanger the generalpublic. Airplanes used in racing and exhibitionflying may be issued NX (experimental)certificates under the terms of this section. Theapplicant shall submit sufficient data, such as

photographs, to identify the airplanesatisfactorily and, upon inspection of theairplane, any pertinent information foundnecessary by the Administrator to safeguard thegeneral public.

(b) An airplane manufactured in accordancewith a type certificate (see §§ 3.15-3.19) andconforming with the type design will becomeeligible for an airworthiness certificate when,upon inspection of the airplane, theAdministrator determines it so to conform andthat the airplane is in a condition for safeoperation. For each newly manufactured airplanethis determination shall include a flight check bythe applicant.

(Amdt. 03-0, 11 F.R. 133.68, as amended byAmdt. 03-4, 13 F.R. 2965)

TYPE CERTIFICATES

§ 3.15 Requirements for issuance. A typecertificate will be issued when the followingrequirements of §§ 3.16 to 3.19 are met.

§ 3.16 Data required for NC and NRcertification. The applicant for a type certificateshall submit to the Administrator suchdescriptive data, test reports, and computationsas are necessary to demonstrate that the airplanecomplies with the airworthiness requirements.The descriptive data shall be known as the typedesign and shall consist of drawings andspecifications disclosing the configuration of theairplane and all design features covered in theairworthiness requirements as well as sufficientinformation on dimensions, materials, andprocesses to define the strength of the structure.The type design shall describe the airplane insufficient detail to permit the airworthiness ofsubsequent airplanes of the same type to bedetermined by comparison with the type design.

[Amdt. 03-0, 11 F.R. 13368, as amended byAmdt. 03-4, 13 F.R. 2965]

§ 3.17 Inspection and tests for NC and NRcertification. The authorized representatives ofthe Administrator shall have access to theairplane and may witness or conduct suchinspections and tests as are necessary todetermine compliance with the airworthinessrequirements.


§ 3.18 Inspection. Inspections and testsshall include all those found necessary by theAdministrator to insure that the airplaneconforms with the following:

(a) All materials and products are inaccordance with the specification given in thetype design.

(b) All parts of the airplane are constructedin accordance with the drawings contained in thetype design.

(c) All manufacturing processes,construction, and assembly are such that thedesign strength and safety contemplated by thetype design will be realized in service.

§ 3.19 Flight tests. (Applicable to allairplanes certficated as a type on or after May 15,1947.) After proof of compliance with thestructural requirements contained in this part,and upon completion of all necessary inspectionand testing on the ground, and proof of theconformity of the airplane with the type design,and upon receipt from the applicant of a report offlight tests conducted by him, there shall beconducted such official flight tests as theAdministrator finds necessary to determinecompliance with §§ 3.61 through 3.780. After theconclusion of these flight tests such additionalflight tests shall be conducted as theAdministrator finds necessary to ascertainwhether there is reasonable assurance that theairplane, its components, and equipment arereliable and function properly. The extent ofsuch additional flight tests shall depend upon thecomplexity of the airplane, the number and natureof new design features, and the record ofprevious tests and experience for the particularairplane model, its components, and equipment.If practicable, the flight tests performed for thepurpose of ascertaining the reliability and properfunctioning shall be conducted on the sameairplane which was used in flight tests to showcompliance with §§ 3.61 through 3.780.


CHANGES

§ 3.23 Changes. Changes shall besubstantiated to demonstrate compliance of theairplane with the appropriate airworthinessrequirements in effect when the particularairplane was certificated as a type, unless the

holder of the type certificate chooses to showcompliance with the currently effectiverequirements subject to the approval of theAdministrator, or unless the Administrator findsit necessary to require compliance with currentairworthiness requirements.

§ 3.24 Minor changes. Minor changes tocertificated airplanes which obviously do notimpair the condition of the airplane for safeoperation shall be approved by the authorizedrepresentatives of the Administrator prior to thesubmittal to the Administrator of any requiredrevised drawings.

§ 3.25 Major changes. A major change isany change not covered by minor changes asdefined in § 3.24.

§ 3.26 Service experience changes. Whenexperience shows that any particular part ofcharacteristic of an airplane is unsafe, the holderof the type certificate for such airplane shallsubmit for approval of the Administrator thedesign changes which are necessary to correctthe unsafe condition after the unsafe conditionbecomes known the Administrator shall withholdthe issuance of airworthiness certificates foradditional airplanes of the type involved until hehas approved the design changes and until theadditional airplanes are modified to include suchchanges. Upon approval by the Administratorthe design changes shall be considered as a partof the type design, and descriptive data coveringthese changes shall be made available by theholder of the type certificate to all owners ofairplanes previously certificated under such typecertificate.

§ 3.27 Application to earlierairworthiness requirements. In the case ofairplanes approved as a type under the terms ofearlier airworthiness requirements, theAdministrator may require that an airplanesubmitted for an original airworthiness certificatecomply with such portions of the currentlyeffective airworthiness requirements as may benecessary for safety.

APPROVAL OF MATERIALS, PARTS, PROCESSES, AND

APPLIANCES

§ 3.31 Specifications. (a) Materials, parts,processes, and appliances shall be approvedupon a basis and in a manner found necessary bythe Administrator to implement the pertinent

provisions of this subchapter. The Administratormay adopt and publish such specifications as hefinds necessary to administer this section, andshall incorporate therein such portions of theaviation industry, Federal, and militaryspecifications respecting such materials, parts,processes, and appliances as he findsappropriate.

(b) Any material, part, process, or applianceshall be deemed to have met the requirements forapproval when it meets the pertinentspecifications adopted by the Administrator, andthe manufacturer so certifies in a mannerprescribed by the Administrator.

[Amdt. 03-3, 12 F.R. 7898]

DEFINITIONS

§ 3.41 Standard atmosphere. Thestandard atmosphere shall be based upon thefollowing assumptions:

(a) The air is a dry perfect gas.

(b) The temperature at sea level is 59° F.

(c) The pressure at sea level is 29.92 inchesHg.

(d) The temperature gradient from sea levelto the altitude at which the temperature becomes-67° F. is -0.003566° F. per foot and zero thereabove.

(e) The density ρo at sea level under theabove conditions is 0.002378 lbs. sec.2/ft4.

§ 3.42 Hot-day condition. See § 3.583.

§ 3.43 Airplane configuration. This termrefers to the position of the various elementsaffecting the aerodynamic characteristics of theairplane, such as landing gear and flaps.

§ 3.44 Weights. Referencesections

Empty weight: The actual weightused as a basis for determiningoperating weights.........…………….. 3.73Maximum weight: The maximumweight at which the airplane mayoperate in accordance with theairworthiness requirements...........…. 3.74Minimum weight: The minimumweight at which compliance with theairworthiness requirements isdemonstrated..................................... 3.75Maximum design weight: The

maximum weight used for thestructural design of the airplane........ 3.181Minimum design weight: Theminimum weight conditioninvestigated in the structural flightload conditions, not greater than theminimum weight specified in §3.75.................................................. 3.181Design landing weight: The weightused in the structural investigation ofthe airplane for normal landingconditions. Under the provisions of §3.242, this weight may be equal to orless than the maximum designweight............................................... 3.242Unit weights for design purposes:Gasoline....................... 6 pounds per United

States gallon.Lubricating oil.............. 7.5 pounds per United

States gallon.Crew and passengers.... 170 pounds per person.

§ 3.45 Power.

One horsepower: 33,000 foot-pounds perminute.

Take-off power: the take-off rating of theengine established in accordance with Part 13,Aircraft Engine Airworthiness.

Maximum continuous power: The maximumcontinuous rating of the engine established inaccordance with Part 13, Aircraft EngineAirworthiness.

§ 3.46 Speeds.

Vt True air speed of the airplane relative tothe undisturbed air.

In the following symbols having subscripts, Vdenotes:

(a) "Equivalent" air speed for structuraldesign purposes equal to Vt√p/po.

(b) "True indicated" or "calibrated" airspeed for performance and operating purposesequal to indicator reading corrected for positionand instrument errors.

Referencesections

Vs0 stalling speed, in the landconfiguration..................................... 3.82Vs1 stalling speed in theconfigurations specified for particularconditions.....................…………….. 3.82

Vsf computed stalling speed atdesign landing weight with flaps fullydefected.....................................

3.190

Vx speed for best angle of climb.Vy speed for best rate of climb.Vmc minimum control speed............. 3.111Vf design speed for flight loadconditions with flaps in landingposition.........................................….

3.190

Vfe flaps-extended speed................... 3.742Vp design maneuvering speed........... 3.184Vc design cruising speed.................. 3.184Vd design dive speed........................ 3.184Vne never-exceed speed.................... 3.739Vno maximum structural cruisingspeed................................................ 3.740Vh maximum speed in level flight atmaximum continuous power.

§ 3.47 Structural terms.

Structure: Those portions of the airplane thefailure of which would seriously endanger thesafety of the airplane.

Design wing area, S: The area enclosed bythe wing outline (including ailerons, and flaps inthe retracted position, but ignoring fillets andfairings) on a surface containing the wing chords.The outline is assumed to extend through thenacelles and fuselage to the centerline ofsymmetry.

Aerodynamic coefficients: CL, CN, CM, etc.,used in this part, are nondimensional coefficientsfor the forces and moments acting on an airfoil,and correspond to those adopted by the UnitedStates National Advisory Committee forAeronautics.

CL = airfoil lift coefficient.

CN = airfoil normal force coefficient (normal towing chord line).

CNA = airplane normal force coefficient (basedon lift of complete airplane and design wing area).

CM = pitching moment coefficient.

LoadsReferenceSections

Limit load: The maximum loadanticipated in service...................…. 8.171Ultimate load: The maximum loadwhich a part of structure must becapable of supporting................….... 8.173Factor of safety: The factor by whichthe limit load must be multiplied to

establish the ultimate load............... 8.172

Load factor or acceleration factor, n: Theratio of the force acting on a mass to the weightof the mass. When the force in questionrepresents the net external load acting on theairplane in a given direction, n represents theacceleration in that direction in terms of thegravitational constant.

Limit load factor: The load factorcorresponding to limit load.

Ultimate load factor: The load factorcorresponding to ultimate load.

§ 3.48 Susceptibility of materials to fire.Where necessary for the purpose of determiningcompliance with any of the definitions in thissection, the Administrator shall prescribe theheat conditions and testing procedures whichany specific material or individual part must meet.

(a) Fireproof. "Fireproof" material means amaterial which will withstand heat equally well orbetter than steel in dimensions appropriate forthe purpose for which it is to be used. Whenapplied to material and parts used to confine firesin designated fire zones "fireproof" means thatthe material or part will perform this functionunder the most severe conditions of fire andduration likely to occur in such zones.

(b) Fire-resistant. When applied to sheetor structural members, "fire-resistant" materialshall mean a material which will withstand heatequally well or better than aluminum alloy indimensions appropriate for the purpose for whichit is to be used. When applied to fluid-carryinglines, this term refers to a line and fittingassembly which will perform its intendedprotective functions under the heat and otherconditions likely to occur at the particularlocation.

(c) Flames-resistant. "Flame-resistant"material means material which will not supportcombustion to the point of propagatting, beyondsafe limits, a flame after removal of the ignitionsource.

(d) Flash-resistant. "Flash-resistant"material means material which will not burnviolently when ignited.

(e) Inflammable. "Inflammable" fluids orgases means those which will ignite readily orexplode.

S UBPART B ——FLIGHT REQUIREMENTS

GENERAL

§ 3.61 Policy re proof of compliance.Compliance with the requirements specified inthis subpart governing functional characteristicsshall be demonstrated by suitable flight or othertests conducted upon an airplane of the type, orby calculations based upon the test data referredto above, provided that the results so obtainedare substantially equal in accuracy to the resultsof direct testing. Compliance with eachrequirement must be provided at the criticalcombination of airplane weight and center ofgravity position within the range of either forwhich certification is desired. Such compliancemust be demonstrated by systematicinvestigation of all probable weight and center ofgravity combinations or must be reasonablyinferable from such as are investigated.

§ 3.62 Flight test pilot. The applicantshall provide a person holding an appropriatepilot certificate to make the flight tests, but adesignated representative of the Administratormay pilot the airplane insofar as that may benecessary for the determination of compliancewith the airworthiness requirements.

§ 3.63 Noncompliance with testrequirements. Official type tests will bediscontinued until corrective measures have beentaken by the applicant when either:

(a) The applicant’s test pilot is unable orunwilling to conduct any of the required flighttests; or

(b) Items of noncompliance withrequirements are found which may renderadditional test data meaningless or are of suchnature as to make further testing undulyhazardous.

§ 3.64 Emergency egress. Adequateprovisions shall be made for emergency egressand use of parachutes by members of the crewduring the flight tests.

§ 3.65 Report. The applicant shall submitto the representative of the Administrator areport covering all computations and testsrequired in connection with calibration ofinstruments used for test purposes andcorrection of test results to standard atmosphericconditions. The representative of theAdministrator will conduct any flight tests which

he finds to be necessary in order to check thecalibration and correction report.

WEIGHT RANGE AND CENTER OF GRAVITY

§ 3.71 Weight and balance. (a) Thereshall be established, as a part of the typeinspection, ranges of weight and center ofgravity within which the airplane may be safelyoperated.

(b) When low fuel adversely affects balanceor stability, the airplane shall be so tested as tosimulate the condition existing when the amountof usable fuel on board does not exceed 1 gallonfor every 12 maximum continuous horsepower ofthe engine or engines installed.

§ 3.72 Use of ballast. Removable ballastmay be used to enable airplanes to comply withthe flight requirements in accordance with thefollowing provisions:

(a) The place or places for carrying ballastshall be properly designed, installed, and plainlymarked as specified in § 3.766.

(b) The Airplane Flight Manual shallinclude instructions regarding the properdisposition of the removable ballast under allloading conditions for which such ballast isnecessary, as specified in § 3.755-3.770.

§ 3.73 Empty weight. The empty weightand corresponding center of gravity locationshall include all fixed ballast, the unusable fuelsupply (see § 3.437), undrainable oil, full enginecoolant, and hydraulic fluid. The weight andlocation of items of equipment installed when theairplane is weighed shall be noted in the AirplaneFlight Manual.

§ 3.74 Maximum weight. The maximumweight shall not exceed any of the following:

(a) The weigh selected by the applicant.

(b) The design weight for which thestructure has been proven.

(c) The maximum weight at whichcompliance with all of the requirements specifiedis demonstrated, and shall not be less than thesum of the weights of the following:

(1) The empty weight as defined by § 3.73.

(2) One gallon of usable fuel (see § 3.437)for every seven maximum continuous horsepowerfor which the airplane is certificated.

(3) The full oil capacity.

(4) 170 pounds in all seats (normalcategory) or 190 pounds in all seats (utility andacrobatic category) unless placarded otherwise.

§ 3.75 Minimum weight. The minimumweight shall not exceed the sum of the weights ofthe following:

(a) The empty weight is defined by § 3.73.

(b) The minimum crew necessary to operatethe airplane (170 pounds for each crew member).

(c) One gallon of usable fuel (see § 3.437)for every 12 maximum continuous horsepower forwhich the airplane is certificated.

(d) Either 1 gallon of oil for each 25 gallonsof fuel specified in (c) or 1 gallon of oil for each75 maximum continuous horsepower for whichthe airplane is certifcated, whichever is greater.

§ 3.76 Center of gravity position. If thecenter of gravity position under any possibleloading condition between the maximum weightas specified in § 3.74 and the minimum weight asspecified in § 3.75 lies beyond (a) the extremesselected by the applicant, or (b) the extremes forwhich the structure has been proven, or (c) theextremes for which compliance with all functionalrequirements were demonstrated, loadinginstructions shall be provided in the AirplaneFlight Manual as specified in § 3.777-3.780.

PERFORMANCE REQUIREMENTS

GENERAL

§ 3.81 Performance. The following itemsof performance shall be determined and theairplane shall comply with the correspondingrequirements in standard atmosphere and still air.

§ 3.82 Definition of stalling speeds. (a)Vso denotes the true indicated stalling speed, ifobtainable, or the minimum steady flight speed atwhich the airplane is controllable, in miles perhour, with:

(1) Engines idling, throttles closed (or notmore than sufficient power for zero thrust),

(2) Propellers in position normally used fortake-off,

(3) Landing gear extended,

(4) Wing flaps in the landing position,

(5) Cowl flaps closed,

(6) Center of gravity in the mostunfavorable position within the allowable landingrange,

(7) The weight of the airplane equal to theweight in connection with which Vso is beingused as a factor to determine a requiredperformance.

(b) Vs1 denotes the true indicated stallingspeed, if obtainable, otherwise the calculatedvalue in miles per hour, with:

(1) Engines idling, throttles closed (or notmore than sufficient power for zero thrust),

(2) Propellers in position normally used fortake-off, the airplane in all other respects (flaps,landing gear, etc.) in the particular conditionexisting in the particular test in connection withwhich Vs1 is being used,

(3) The weight of the airplane equal to theweight in connection with which Vs1 is beingused as a factor to determine a requiredperformance.

(c) These speeds shall be determined byflight tests using the procedure outlined in§3.120.

§ 3.83 Stalling speed. Vso at maximumweight shall not exceed 70 miles per hour for (1)single-engine airplanes and (2) multiengineairplanes which do not have the rate of climb withcritical engine inoperative specified in §3.85 (b).

TAKE-OFF

§ 3.84 Take-off. (a) The distance requiredto take off and climb over a 50-foot obstacle shallbe determined under the following conditions:

(1) Most unfavorable combination ofweight and center of gravity location,

(2) Engines operating within the approvedlimitations,

(3) Cowl flaps in the position normally usedfor take-off.

(b) Upon obtaining a height of 50 feetabove the level take-off surface, the airplane shallhave attained a speed of not less than 1.3 Vs1

unless a lower speed of not less than Vx plus 5can be shown to be safe under all conditions,including turbulence and complete engine failure.

(c) The distance so obtained, the type ofsurface from which made, and the pertinentinformation with respect to the cowl flap position,the use of flight-path control devices and landinggear retraction system shall be entered in theAirplane Flight Manual. The take-off shall bemade in such a manner that its reproduction shallnot require an exceptional degree of skill on thepart of the pilot or exceptionally favorableconditions.

CLIMB

§ 3.85 Climb—(a) Normal climbcondition. The steady rate of climb at sea levelshall be at least 300 feet per minute, and thesteady angle of climb at least 1:12 for landplanesor 1:15 for seaplanes with:

(1) Not more than maximum continuouspower on all engines,

(2) Landing gear fully retracted,

(3) Wing flaps in take-off position,

(4) Cowl flaps in the position used incooling tests specified in §§ 3.581-3.596.

(b) Climb with inoperative engine. Allmuliengine airplanes having a stalling speed Vsogreater than 70 miles per hour or a maximumweight greater than 6,000 pounds shall have asteady rate of climb of at least 0.02 Vso in feet perminute at an altitude of 5,000 feet with the criticalengine inoperative and:

(1) The remaining engines operating at notmore than maximum continuous power,

(2) The inoperative propeller in theminimum drag position,

(3) Landing gear retracted,

(4) Wing flaps in the most favorableposition,

(5) Cowl flaps in the position used incooling tests specified in §§ 3.581-3.596.

(c) Balked landing conditions. The steadyangle of climb at sea level shall be at least 1:30with:

(1) Take-off power on all engines,

(2) Landing gear extended,

(3) Wing flaps in landing position.

If rapid retraction is possible with safetywithout loss of altitude and without requiringsudden changes of angle of attack or exceptionalskill on the part of the pilot, wing flaps may beretracted.

LANDING

§ 3.86 Landing (a) The horizontaldistance required to land and to come to acomplete stop (to a speed of approximately 3miles per hour for seaplanes or float planes) froma point at a height of 50 feet above the landingsurface shall be determined as follows:

(1) Immediately prior to reaching the 50-footaltitude, a steady gliding approach shall havebeen maintained, with a true indicated air speedof at least 1.3 Vso.

(2) The landing shall be made in such amanner that there is no excessive verticalacceleration, no tendency to bounce, nose over,ground loop, porpoise, or water loop, and in sucha manner that its reproduction shall not requireany exceptional degree of skill on the part of thepilot or exceptionally favorable conditions.

(b) The distance so obtained, the type oflanding surface on which made and the pertinentinformation with respect of cowl flap position,and the use of flight path control devices shall beentered in the Airplane Flight Manual.

FLIGHT CHARACTERISTICS

§ 3.105 Requirements. The airplane shallmeet the requirements set forth in §§ 3.106 to3.124 at all normally expected operating altitudesunder all critical loading conditions within therange of center of gravity and, except asotherwise specified, at the maximum weight forwhich certification is sought.

CONTROLLABILITY

§ 3.106 General. The airplane shall besatisfactorily controllable and maneuverableduring take-off, climb, level flight, drive, andlanding with or without power. It shall bepossible to make a smooth transition from oneflight condition to another, including turns andslips, without requiring an exceptional degree ofskill, alertness, or strength on the part of thepilot, and without danger of exceeding the limitload factor under all conditions of operationprobable for the type, including for multiengineairplanes those conditions normally encountered

in the event of sudden failure of any engine.Compliance with "strength of pilots" limits neednot be demonstrated by quantitative tests unlessthe Administrator finds the condition to bemarginal. In the latter case they shall not exceedmaximum values found by the Administrator tobe appropriate for the type but in no case shallthey exceed the following limits:

Pitch Roll Yaw(a) For temporaryapplication: Stick........................... Wheel 1......................

6075

3060

150150

(b) For prolongedapplication...................... 10 5 201Applied to rim.

§ 3.107-U Approved acrobatic maneuvers. Itshall be demonstrated that the approvedacrobatic maneuvers can be performed safely.Safe entry speeds shall be determined for thesemaneuvers.

§ 3.108-A Acrobatic maneuvers. It shall bedemonstrated that acrobatic maneuvers can beperformed readily and safely. Safe entry speedsshall be determined for these maneuvers.

§ 3.109 Longitudinal control. The airplaneshall be demonstrated to comply with thefollowing requirements:

(a) It shall be possible at all speeds belowVx to pitch the nose downward so that the rate ofincrease in air speed is satisfactory for promptacceleration of Vx with:

(1) Maximum continuous power on allengines, the airplane trimmed at Vx.

(2) Power off, the airplane trimmed at 1.4Vs1.

(3) (i) Wing flaps and landing gearextended and

(ii) Wing flaps and landing gear retracted.

(b) During each of the controllabilitydemonstrations outlined below it shall not requirea change in the trim control or the exertion ofmore control force than can be readily appliedwith one hand for a short period. Each maneuvershall be performed with the landing gearextended.

(1) With power off, flaps retracted, and theairplane trimmed at 1.4 Vs1, the flaps shall be

extended as rapidly as possible while maintainingthe air speed at approximately 40 percent abovethe instantaneous value of the stalling speed.

(2) Same as subparagraph (1) of thisparagraph, except the flaps shall be initiallyextended and the airplane trimmed at 1.4 Vs1, thenthe flaps shall be retracted as rapidly as possible.

(3) Same as subparagraph (2) of thisparagraph, except maximum continuous powershall be used.

(4) With power off, the flaps retracted, andthe airplane trimmed at 1.4 Vs1, take-off powershall be applied quickly while the same air speedis maintained.

(5) Same as subparagraph (4) of thisparagraph, except with the flaps extended.

(6) With power off, flaps extended, and theairplane trimmed at 1.4 Vs1, air speeds within therange of 1.1 Vs1 to 1.7 Vs1 or Vf whichever is thelesser, shall be obtained and maintained.

(c) It shall be possible without the use ofexceptional piloting skill to maintain essentiallylevel flight when flap retraction from any positionis initiated during steady horizontal flight at 1.1Vs1 with simultaneous application of not morethan maximum continuous power.

§ 3.110 Lateral and directional control.(a) It shall be possible with multiengine airplanesto execute 15-degree banked turns both with andagainst the inoperative engine from steady climbat 1.4 Vs1 or Vy for the condition with:

(1) Maximum continuous power on theoperating engines,

(2) Rearmost center of gravity,

(3) (i) Landing gear retracted and (ii)Landing gear extended.

(4) Wing flaps in most favorable climbposition,

(5) Maximum weight,

(6) The inoperative propeller in its minimumdrag condition.

(b) It shall be possible with multiengineairplanes, while holding the wings level laterallywithin 5 degrees, to execute sudden changes inheading in both directions without dangerouscharacteristics being encountered. This shall bedemonstrated at 1.4 Vs1 or Vy up to heading

changes of 15 degrees, except that the headingchange at which the rudder force corresponds tothat specified in § 3.106 need not be exceeded,with:

(1) The critical engine inoperative,

(2) Maximum continuous power on theoperating engine(s),

(3) (i) Landing gear retracted and (ii)Landing gear extended,

(4) Wing flaps in the most favorable climbposition,

(5) The inoperative propeller in its minimumdrag condition,

(6) The airplane center of gravity at itsrearmost position.

§ 3.111 Minimum control speed (Vmc). (a)A minimum speed shall be determined under theconditions specified below, such that when anyone engine is suddenly made inoperative at thatspeed, it shall be possible to recover control ofthe airplane, with the one engine still inoperative,and to maintain it in straight flight at that speed,either with zero yaw or, at the option of theapplicant, with a bank not in excess of 5 degrees.Such speed shall not exceed 1.3 Vs1, with:

(1) Take-off or maximum available power onall engines,

(2) Rearmost center of gravity,

(3) Flaps in take-off position,

(4) Landing gear retracted.

(b) In demonstrating this minimum speed,the rudder force required to maintain it shall notexceed forces specified in § 3.106, nor shall it benecessary to throttle the remaining engines.During recovery the airplane shall not assumeany dangerous attitude, nor shall it requireexceptional skill, strength, or alertness on the partof the pilot to prevent a change of heading inexcess of 20 degrees before recovery is complete.

TRIM

§ 3.112 Requirements. (a) The means usedfor trimming the airplane shall be such that, afterbeing trimmed and without further pressure upon

or movement of either the primary control or itscorresponding trim control by the pilot or theautomatic pilot, the airplane will maintain:

(1) Lateral and directional trim in level flightat a speed of 0.9 Vh or at Vc, if lower, with thelanding gear and wing flaps retracted:

(2) Longitudinal trim under the followingconditions:

(i) During a climb with maximumcontinuous power at a speed between Vx and 1.4Vs1,

(a) With landing gear retracted and wingflaps retracted,

(b) With landing gear retracted and wingflaps in the take-off position.

(ii) During a glide with power off at a speednot in excess of 1.4 Vs1,

(a) With landing gear extended and wingflaps retracted,

(b) With landing gear extended and wingflaps extended under the forward center ofgravity position approved with the maximumauthorized weight.

(c) With landing gear extended and wingflaps extended under the most forward center ofgravity position approved, regardless of weight.

(iii) During level flight at any speed from 0.9Vh to Vx or 1.4 Vs1 with landing gear and wingflaps retracted.

(b) In addition to the above, multiengineairplanes shall maintain longitudinal anddirectional trim at a speed between Vy and 1.4 Vs1

during climbing flight with the critical of two ormore engines inoperative, with:

(1) The other engine(s) operating atmaximum continuous power.

(2) The landing gear retracted,

(3) Wing flaps retracted,

(4) Bank not in excess of 5 degrees.

S TABILITY

§ 3.113 General. The airplane shall belongitudinally, directionally, and laterally stablein accordance with the following sections.Suitable stability and control "feel" (staticstability) shall be required in other conditions

normally encountered in service, if flight testsshow such stability to be necessary for safeoperation.

§ 3.114 Static longitudinal stability. In theconfigurations outlined in § 3.115 and with theairplane trimmed as indicated, the characteristicsof the elevator control forces and the frictionwithin the control system shall be such that:

(a) A pull shall be required to obtain andmaintain speeds below the specified trim speedand a push to obtain and maintain speeds abovethe specified trim speed. This shall be so at anyspeed which can be obtained without excessivecontrol force, except that such speeds need notbe greater than the appropriate maximumpermissible speed or less than the minimumspeed in steady unstalled flight.

(b) The air speed shall return to within 10percent of the original trim speed when thecontrol force is slowly released from any speedwithin the limits defined in paragraph (a) of thissection.

§ 3.115 Specific conditions. In conditionsset forth in this section, within the speedsspecified, the stable slope of stick force versusspeed curve shall be such that nay substantialchange in speed is clearly perceptible to the pilotthrough a resulting change in stick force.

(a) Landing. The stick force curve shallhave a stable slope and the stick force shall notexceed 40 lbs. at any speed between 1.1 Vs1 and1.3 Vs1 with:

(1) Wing flaps in the landing position,

(2) The landing gear extended,

(3) Maximum weight,

(4) Throttles closed on all engines,

(5) The airplane trimmed at 1.4 Vs1 withthrottles closed.

(b) Climb. The stick force curve shall havea stable slope at all speeds between 1.2 Vs1 and1.6 Vs1 with:

(1) Wing flaps retracted,

(2) Landing gear retracted,

(3) Maximum weight,

(4) 75 percent of maximum continuouspower,

(5) The airplane trimmed at 1.4 Vs1.

(c) Cruising. (1) Between 1.3 Vs1 and themaximum permissible speed, the stick force curveshall have a stable slope at all speeds obtainablewith a stick force not in excess of 40 pounds with:

(i) Landing gear retracted,

(ii) Wing flaps retracted,

(iii) Maximum weight,

(iv) 75 percent of maximum continuouspower,

(v) The airplane trimmed for level flight with75 percent of the maximum continuous power.

(2) Same as subparagraph (1) of thisparagraph, except that the landing gear shall beextended and the level flight trim speed need notbe exceeded.

§ 3.116 Instrumented stick forcemeasurements. Instrumented stick forcemeasurements need not be made when changesin speed are clearly reflected by changes in stickforces and the maximum forces obtained in theabove conditions are not excessive.

§ 3.117 Dynamic longitudinal stability.Any short period oscillation occurring betweenstalling speed and maximum permissible speedshall be heavily damped with the primary controls(1) free, and (2) in a fixed position.

§ 3.118 Directional and lateral stability—(a) Three-control airplanes. (1) The staticdirectional stability, as shown by the tendency torecover from a skid with rudder free, shall bepositive for all flap positions and symmetricalpower conditions, and for all speeds from 1.2 Vs1

up to the maximum permissible speed.

(2) The static lateral stability as shown bythe tendency to raise the low wing in a sideslip,for all flap positions and symmetrical powerconditions, shall:

(i) Be positive at the maximum permissiblespeed.

(ii) Not be negative at a speed equal to 1.2Vs1.

(3) In straight steady sideslips(unaccelerated forward slips), the aileron andrudder control movements and forces shallincrease steadily, but not necessarily in constantproportion, as the angle of sideslip is increased;

the rate of increase of the movements and forcesshall lie between satisfactory limits up to sideslipangles considered appropriate to the operation ofthe type. At greater angles, up to that at whichthe full rudder control is employed or a rudderpedal force of 150 pounds is obtained, the rudderpedal forces shall not reverse and increasedrudder deflection shall produce increased anglesof sidelslip. Sufficient bank shall accompanysideslipping to indicate adequately any departurefrom steady unyawed flight.

(4) Any short-period oscillation occurringbetween stalling speed and maximum permissiblespeed shall be heavily damped with the primarycontrols (i) free and (ii) in a fixed position.

(b) Two-control (or simplified) airplanes.(1) The directional stability shall be shown to beadequate by demonstrating that the airplane in allconfigurations can be rapidly rolled from a 45-degree bank to a 45-degree bank in the oppositedirection without exhibiting dangerous skiddingcharacteristics.

(2) Lateral stability shall be shown to beadequate by demonstrating that the airplane willnot assume a dangerous attitude or speed whenall the controls are abandoned for a period of 2minutes. This demonstration shall be made inmoderately smooth air with the airplane trimmedfor straight level flight at 0.9 Vh (or at Vc, iflower), flaps and gear retracted, and withrearward center of gravity loading.

(3) Any short period oscillation occurringbetween the stalling speed and the maximumpermissible speed shall be heavily damped withthe primary controls (i) free and (ii) in a fixedposition.

S TALLS

§ 3.120 Stalling demonstration. (a) Stallsshall be demonstrated under two conditions:

(1) With power off.

(2) With the power setting not less thanthat required to show compliance with § 3.85 (a).

(b) In either condition it shall be possible,with flaps and landing gear in any position, withcenter of gravity in the position least favorablefor recovery, and with appropriate airplaneweights for: (1) Airplanes having independentlycontrolled rolling and directional controls toproduce and to correct roll by unreversed use of

the rolling control and to produce and to correctyaw by unreversed use of the directional controlduring the maneuvers described below up to thetime when the airplane pitches, (2) two-controlairplanes having either interconnected lateral anddirectional controls or providing only one ofthese controls to produce and to correct roll byunreversed use of the rolling control withoutproducing excessive yaw during the maneuversdescribed below up to the time the airplanepitches.

(c) During the recovery portions of themaneuver, pitch shall not exceed 30 degreesbelow level, there shall be no loss of altitude inexcess of 100 feet, and not more than 15 degreesroll or yaw shall occur when controls are notused for 1 second after pitch starts and are usedthereafter only in a normal manner.

(d) Where clear and distinctive stallwarning is apparent to the pilot at a speed at least5 percent above the stalling speed with flaps andlanding gear in any position, both in straight andturning flight, these requirements are modified asfollows:

(1) It shall be possible to prevent more than15 degrees roll or yaw by the normal use ofcontrols.

(2) Any loss of altitude in excess of 100 feetor any pitch in excess of 30 degrees below levelshall be entered in the Airplane Flight Manual.

(e) In demonstrating the qualities set forthin paragraph (d) of this section, the order ofevents shall be:

(1) With trim controls adjusted for straightflight at a speed of approximately 1.4 Vs1, reducespeed by means of the elevator control until thespeed is steady at slightly above stalling speed,then

(2) Pull elevator control back at a rate suchthat the airplane speed reduction does not exceed1 miler per hour per second until a stall isproduced as evidenced by an uncontrollabledownward pitching motion of the airplane, oruntil the control reaches the stop. Normal use ofthe elevator control for recovery may be madeafter such pitching motion is unmistakablydeveloped.

§ 3.121 Climbing stalls. When stalled froman excessive climb attitude it shall be possible torecover from this maneuver without exceeding

the limiting air speed or the allowable accelerationlimit.

§ 3.122 Turning flight stalls. When stalledduring a coordinated 30-degree banked turn with75 percent maximum continuous power on allengines, flaps and landing gear retracted, it shallbe possible to recover to normal level flightwithout encountering excessive loss of altitude,uncontrollable rolling characteristics, oruncontrollable spinning tendencies. Thesequalities shall be demonstrated by performing thefollowing maneuver: After a steady curvilinearlevel coordinated flight condition in a 30-degreebank is established and while maintaining the 30-degree bank, the airplane shall be stalled bysteadily and progressively tightening the turnwith the elevator control until the airplane isstalled or until the elevator has reached its stop.When the stall has fully developed, recovery tolevel flight shall be made with normal use of thecontrols.

§ 3.123 One-engine-inoperative stalls.Multiengine airplanes shall not display anyundue spinning tendency and shall be safelyrecoverable without applying power to theinoperative engine when stalled with:

(a) The critical engine inoperative,

(b) Flaps and landing gear retracted,

(c) The remaining engines operating at upto 75 percent of maximum continuous power,except that the power need not be greater thanthat at which the use of maximum control traveljust holds the wings laterally level in approachingthe stall. The operating engines may be throttledback during the recovery from the stall.

S PINNING

§ 3.124 Spinning—(a) Category N. Allairplanes of 4,000 pounds or less maximum weightshall recover from a one-turn spin with controlsassisted to the extent necessary to overcomefriction in not more than one and one-halfadditional turns and without exceeding either thelimiting air speed or the limit positivemaneuvering load factor for the airplane. It shallnot be possible to obtain uncontrollable spins bymeans of any possible use of the controls.Compliance with the above shall be demonstratedat any permissible combination of weight andcenter of gravity positions obtainable with all orpart of the design useful load. All airplanes in

this category, regardless of weight, shall beplacarded against spins or demonstrated to be"characteristically incapable of spinning" inwhich case they shall be so designated. (Seeparagraph (d) of this section.)

(b) Category U. Airplanes in this categoryshall comply with either the entire requirementsof paragraph (a) of this section or the entirerequirements of paragraph (c) of this section.

(c) Category A. All airplanes in thiscategory must be capable of spinning and shallcomply with the following:

(1) At any permissible combination ofweight and center of gravity position obtainablewith all or part of the design useful load, theairplane shall recover from a six-turn spin withcontrols free in not more than four additionalturns after releasing the controls. If the airplanewill not recover as prescribed with controls freebut will recover with the controls assisted to theextent necessary to overcome friction, theairplane may be certificated with the rearmostcenter of gravity position 2 percent forward ofthe position used in the test.

(2) It shall be possible to recover at anypoint in the spinning described above by usingthe controls in a normal manner for that purposein not more than one and one-half additionalturns, and without exceeding either the limitingair speed or the limit positive maneuvering loadfactor for the airplane. It shall not be possible toobtain uncontrollable spins by means of anypossible use of the controls.

(d) Category NU. When it is desired todesignate an airplane as a type "characteristicallyincapable of spinning," the flight tests todemonstrate this characteristic shall also beconducted with:

(1) A maximum weight 5 percent in excessof the weight for which approval is desired,

(2) A center of gravity at least 3 percent aftof the rearmost position for which approvals isdesired,

(3) An available up-elevator travel 4degrees in excess of that to which the elevatortravel is to be limited by appropriate stops.

(4) An available rudder travel 7 degrees, inboth directions, in excess of that to which the

rudder travel is to be limited by appropriatestops.

GROUND AND WATER CHARACTERISTICS

§ 3.143 Requirements. All airplanes shallcomply with the requirements of §§ 3.144 to3.147.

§ 3.144 Longitudinal stability and control.There shall be no uncontrollable tendency forlandplanes to nose over in any operatingcondition reasonably expected for the type, orwhen rebound occurs during landing or take-off.Wheel brakes shall operate smoothly and shallexhibit no undue tendency to induce nosingover. Seaplanes shall exhibit no dangerous oruncontrollable proposing at any speed at whichthe airplane is normally operated on the water.

§ 3.145 Directional stability and control.(a) There shall be no uncontrollable loopingtendency in 90-degree cross winds up to avelocity equal to 0.2 Vso at any speed at whichthe aircraft may be expected to be operated uponthe ground or water.

(b) All landplanes shall be demonstrated tobe satisfactorily controllable with no exceptionaldegree of skill or alternates on the part of thepilot in power-off landings at normal landingspeed and during which brakes or engine powerare not to maintain a straight path.

(c) Means shall be provided for adequatedirectional control during taxiing.

§ 3.146 Shock absorption. The shockabsorbing mechanism shall not produce damageto the structure when the airplane is taxied on theroughest ground which it is reasonable to expectthe airplane to encounter in normal operation.

§ 3.147 Spray characteristics. Forseaplanes, spray during taxiing, take-off, andlanding shall at no time dangerously obscure thevision of the pilots nor produce damage to thepropeller or other parts of the airplane.

FLUTTER AND VIBRATION

§ 3.159 Flutter and vibration. All parts ofthe airplane shall be demonstrated to be free fromflutter and excessive vibration under all speedand power conditions appropriate to theoperation of the airplane up to at least theminimum valve permitted for Vd in § 3.184. Thereshall also be no buffeting condition in any normal

flight condition severe enough to interfere withthe satisfactory control of the airplane or tocause excessive fatigue to the crew or result instructural damage. However, buffeting as stallwarning is considered desirable anddiscouragement of this type of buffeting is notintended.

S UBPART C——S TRENGTH REQUIREMENTS

GENERAL

§ 3.171 Loads. (a) Strength requirementsare specified in terms of limit and ultimate loads.Limit loads are the maximum loads anticipated inservice. Ultimate loads are equal to the limitloads multiplied by the factor of safety. Unlessotherwise described, loads specified are limitloads.

(b) Unless otherwise provided, thespecified air, ground, and water loads shall beplaced in equilibrium with inertia forces,considering all items of mass in the airplane. Allsuch loads shall be distributed in a mannerconservatively approximating or closelyrepresenting actual conditions. If deflectionsunder load would change significantly thedistribution of external or internal loads, suchredistribution shall be taken into account.

§ 3.172 Factor of safety. The factor ofsafety shall be 1.5 unless otherwise specified.

§ 3.173 Strength and deformations. Thestructure shall be capable of supporting limitloads without suffering detrimental permanentdeformations. At all loads up to limit loads, thedeformation shall be such as not to interfere withsafe operation of the airplane. The structure shallbe capable of supporting ultimate loads withoutfailure for at least 3 seconds, except that whenproof of strength is demonstrated by dynamictests simulating actual conditions of loadapplication, the 3-second limit does not apply.

§ 3.174 Proof of structure. Proof ofcompliance of the structure with the strength anddeformation requirements of § 3.173 shall bemade for all critical loading conditions. Proof ofcompliance by means of structural analysis willbe accepted only when the structure conformswith types for which experience has shown suchmethods to be reliable. In all other casessubstantiating load tests are required. In allcases certain portions of the structure must besubjected to tests as specified in Subpart D.

FLIGHT LOADS

§ 3.181 General. Flight load requirementsshall be complied with at critical altitudes withinthe range in which the airplane may be expectedto operate and at all weights between theminimum design weight and the maximum designweight, with any practicable distribution ofdisposable load within prescribed operatinglimitations stated in § 3.777-3.780.

§ 3.182 Definition of flight load factor.The flight load factors specified represent theacceleration component (in terms of thegravitational constant g) normal to the assumedlongitudinal axis of the airplane, and equal inmagnitude and opposite in direction to theairplane inertia load factor at the center ofgravity.

S YMMETRICAL FLIGHT CONDITIONS (FLAPS

RETRACTED)

§ 3.183 General. The strengthrequirements shall be met at all combinations ofair speed and load factor on and within theboundaries of a pertinent V-n diagram,constructed similarly to the one shown in Figure3-1, which represents the envelope of the flightloading conditions specified by the maneuveringand gust criteria of §§ 3.185 and 3.187. Thisdiagram will also be used in determining theairplane structural operating limitations asspecified in Subpart G.

§ 3.184 Design air speeds. The design airspeeds shall be chosen by the designer exceptthat they shall not be less than the followingvalues:

Vc (design cruising speed)

= 38 √ W/S (NU)= 42 √ W/S (A)

except that for values of W/S greater than 20, theabove numerical multiplying factors shall bedecreased linearly with W/S to a value of 33 atW/S=100: And further provided, That therequired minimum value need be no greater than0.9 Vh actually obtained at sea level.

Vd (design dive speed)=1.40 Vc min (N)=1.50 Vc min (U)=1.55 Vc min (A)

except that for values of W/S greater than 20, theabove numerical multiplying factors shall bedecreased linearly with W/S to a value of 1.35 at

W/S=100. (Vc min is the required minimum valueof design cruising speed specified above.)

Vp (design maneuvering speed)

= Vs √n where:Vs =a computed stalling speed with flaps fully

retracted at the design weight, normallybased on the maximum airplane normalforce coefficient, CNA.

N= limit maneuvering load factor used indesign,

except that the value of Vp need not exceed thevalue of Vc used in design.

§ 3.185 Maneuvering envelope. Theairplane shall be assumed to subjected tosymmetrical maneuvers resulting in the followinglimit load factors, except where limited bymaximum (static) lift coefficients:

(a) The positive maneuvering load factorspecified in § 3.186 at all speeds up to Vd,

(b) The negative maneuvering load factorspecified in § 3.188 at speed Vc; and factorsvarying linearly with speed from the specifiedvalue at Vc to 0.0 at Vd for the N category and -1.0at Vd for the A and U categories.

§ 3.186 Maneuvering load factors. (a) Thepositive limit maneuvering load factors shall notbe less than the following values (see Fig. 3-2):

n = 2.1+ 24,000 --------------Category NW + 10,000

except that n need not be greater than 3.8 andshall not be less than 2.5. For airplanescertificated as characteristically incapable ofspinning, n need not exceed 3.5.

n = 4.4--------------------------------Category Un = 6.0--------------------------------Category A

(b) The negative limit maneuvering loadfactors shall not be less than -0.4 times thepositive load factor for the N and U categories,and shall not be less than -0.5 times the positiveload factor for the A category.

(c) Lower values of maneuvering loadfactor may be employed only if it be proven thatthe airplane embodies features of design whichmake it impossible to exceed such values in flight.(See also § 3.106.)

§ 3.187 Gust envelope. The airplane shallbe assumed to encounter symmetrical verticalgusts as specified below while in level flight andthe resulting loads shall be considered limitloads:

(a) Positive (up) and negative (down) gustsof 30 feet per second nominal intensity at allspeeds up to Vc,

(b) Positive and negative 15 feet per secondgusts at Vd. Gust load factors shall be assumedto vary linearly between Vc and Vd.

§ 3.188 Gust load factors. In applying thegust requirements, the gust load factors shall becomputed by the following formula:

n = 1 + KUVm

575 (W/S)

where: K = ½ (W/S) ¼ (for W/S < 16 p.s.f.)

= 1.33 - 2.67 (for W/S > 16 p.s.f.)(W/S) ¼

U = nominal gust velocity, f.p.s.(Note that the "effective sharp-edged gust"

equals KU.)V = airplane speed, m.p.h.m = slope of lift curve, CL per radian,

corrected for aspect ratio.W/S = wing loading, p.s.f.

§ 3.189 Airplane equilibrium. Indetermining the wing loads and linear inertialoads corresponding to any of the abovespecified flight conditions, the appropriatebalancing horizontal tail load (see § 3.215) shallbe taken into account in a rational orconservative manner.

Incremental horizontal tail loads due tomaneuvering and gusts (see §§ 3.216 and 3.217)shall be reacted by angular inertia of the completeairplane in a rational or conservative manner.

FLAPS EXTENDED FLIGHT CONDITIONS

§ 3.190 Flaps extended flight conditions.(a) When flaps or similar high lift devicesintended for use at the relatively low air speedsof approach, landing, and take-off are installed,the airplane shall be assumed to be subjected tosymmetrical maneuvers and gusts with the flapsfully deflected at the design flap speed Vfresulting in limit load factors within the rangedetermined by the following conditions:

(1) Maneuvering, to a positive limit loadfactor of 2.0.

(2) Positive and negative 15-feet-per-second gusts acting normal to the flight path inlevel flight. The gust load factors shall becomputed by the formula of § 3.188.

Vf shall be assumed not less than 1.4 Vs of 1.8Vsf whichever is greater, where:

Vs = the computed stalling speed with flapsfully retracted at the design weight

Vsf = the computed stalling speed with flapsfully extended at the design weight

except that when an automatic flap loadlimiting device is employed, the airplane may bedesigned for critical combinations of air speedand flap position permitted by the device. (Seealso § 3.338.)

(b) In designing the flaps and supportingstructure, slipstream effects shall be taken intoaccount as specified in § 3.223.

Note: In determining the external loads on theairplane as a whole, the thrust, slip-stream, andpitching acceleration may be assumed equal tozero.

UNSYMMETRICAL FLIGHT CONDITIONS

§ 3.191 Unsymmetrical flight conditions.The airplane shall be assumed to be subjected torolling and yawing maneuvers as described in thefollowing conditions. Unbalanced aerodynamicmoments about the center of gravity shall bereacted in a rational or conservative mannerconsidering the principal masses furnishing thereacting inertia forces.

(a) Rolling conditions. The airplane shallbe designed for (1) unsymmetrical wing loadsappropriate to the category, and (2) the loadsresulting from the aileron deflections and speedsspecified in § 3.222, in combination with anairplane load factor of at least two-thirds of thepositive maneuvering factor used in the design ofthe airplane. Only the wing and wing bracingneed be investigated for this condition.

Note: These conditions may be covered asnoted below:

(a) Rolling accelerations may be obtainedby modifying the symmetrical flight conditionsshown in Figure 3-1 as follows:

(1) Acrobatic category. In conditions Aand F assume 100 percent of the wing air loadacting on one side of the plane of symmetry and60 percent on the other.

(2) Normal and utility categories. Incondition A, assume 100 percent of the wing airload acting on one side of the airplane and 70percent on the other. For airplanes over 1,000pounds design weight, the latter percentage may

be increased linearly with weight up to 80 percentat 25,000 pounds.

(b) The effect of aileron displacement onwing torsion may be accounted for by adding thefollowing increment to the basic airfoil momentcoefficient over the aileron portion of the span inthe critical condition as determined by the noteunder § 3.222:

cm = -.01δ

where:

cm = moment coefficient increment

δ = down aileron deflection in degrees incritical condition

(b) Yawing conditions. The airplane shallbe designed for the yawing loads resulting fromthe vertical surface loads specified in §§ 3.219 to3.221.

S UPPLEMENTARY CONDITIONS

§ 3.194 Special condition for rear lift truss.When a rear lift truss is employed, it shall bedesigned for conditions of reversed airflow at adesign speed of:

V = 10 √ W/S + 10 (m.p.h.)

Note: It may be assumed that the value of CL

is equal to -0.8 and the chordwise distribution istriangular between a peak at the trailing edge andzero at the leading edge.

§ 3.195 Engine torque effects. (a) Enginemounts and their supporting structures shall bedesigned for engine torque effects combined withcertain basic flight conditions as described insubparagraphs (1) and (2) of this paragraph.Engine torque may be neglected in the otherflight conditions.

(1) The limit torque corresponding to take-off power and propeller speed actingsimultaneously with 75 percent of the limit loadsfrom flight condition A. (See Fig. 3-1.)

(2) The limit torque corresponding tomaximum continuous power and propeller speed,acting simultaneously with the limit loads fromflight condition A. (See Fig. 3-1.)

(b) The limit torque shall be obtained bymultiplying the mean torque by a factor of 1.33 inthe case of engines having 5 or more cylinders.For 4-, 3-, and 2-cylinder engines, the factor shallbe 2, 3, and 4, respectively.

§ 3.196 Side load on engine mount. Thelimit load factor in a lateral direction for thiscondition shall be at least equal to one-third ofthe limit load factor for flight condition A (seeFig. 3-1) except that it shall not be less than 1.33.Engine mounts and their supporting structureshall be designed for this condition which may beassumed independent of other flight conditions.

CONTROL S URFACE LOADS

§ 3.211 General. The control surface loadsspecified in the following sections shall beassumed to occur in the symmetrical andunsymmetrical flight conditions as described in§§ 3.189-3.191. See Figures 3-3 to 3-10 foracceptable values of control surface loadingswhich are considered as conforming to thefollowing detailed rational requirements.

§ 3.212 Pilot effort. In the control surfaceloading conditions described, the airloads on themovable surfaces and the correspondingdeflections need not exceed those which couldbe obtained in flight by employing the maximumpilot control forces specified in Figure 3-11. Inapplying this criterion, proper consideration shallbe given to the effects of control system boostand servo mechanisms, tabs, and automatic pilotsystems in assisting the pilot.

§ 3.213 Trim tab effects. The effects of trimtabs on the control surface design conditionsneed be taken into account only in cases wherethe surface loads are limited on the basis ofmaximum pilot effort. In such cases the tabs shallbe considered to be deflected in the directionwhich would assist the pilot and the deflectionshall correspond to the maximum expected degreeof "out of trim" at the speed for the conditionunder consideration.

HORIZONTAL TAIL S URFACES

§ 3.214 Horizontal tail surfaces. Thehorizontal tail surfaces shall be designed for theconditions set forth in §§ 3.215-3.218.

§ 3.215 Balancing loads. A horizontal tailbalancing load is defined as that necessary tomaintain the airplane in equilibrium in a specifiedflight condition with zero pitching acceleration.The horizontal tail surfaces shall be designed forthe balancing loads occurring at any point on thelimit maneuvering envelope, Figure 3-1, and in theflap conditions. (See § 3.190.)

Note: The distribution of Figure 3-7 may beused.

§ 3.216 Maneuvering loads. (a) Atmaneuvering speed Vp assume a suddendeflection of the elevator control to the maximumupward deflection as limited by the control stopsor pilot effort, whichever is critical.

Note: The average loading of Figure 3-3 andthe distribution of Figure 3-8 may be used. Indetermining the resultant normal force coefficientfor the tail under these conditions, it will bepermissible to assume that the angle of attack ofthe stabilizer with respect to the resultantdirection of air flow is equal to that which occurswhen the airplane is in steady unacceleratedflight at a flight speed equal to Vp. The maximumelevator deflection can then be determined fromthe above criteria and the tail normal forcecoefficient can be obtained from the data given inNACA Report No. 688, "AerodynamicCharacteristics of Horizontal Tail Surfaces," orother applicable NACA reports.

(b) Same as case (a) except that the elevatordeflection is downward.

Note: The average loading of Figure 3-3 andthe distribution of Figure 3-8 may be used.

(c) At all speeds above Vp the horizontal tailshall be designed for the maneuveringloads resulting from a sudden upwarddeflection of the elevator, followed by adownair deflection of the elevator suchthat the following combinations ofnormal acceleration and angularacceleration are obtained:

ConditionAirplanenormal

acceleration n

Angularaccelerationradian/sec.2

Down load --- 1.0

Up load ------- nm

Acceptable values of limit averagemaneuvering control surface loadings can beobtained from Figure 3-3 (b) as follows:

HORIZONTAL TAIL S URFACES

(1) Condition § 3.216 (a):Obtain as function of W/S and surface

deflection;Use Curve C for deflection 10° or less;Use Curve B for deflection 20°;

Use Curve A for deflection 30° or more;(Interpolate for other deflections);Use distribution of Figure 3-8.

(2) Condition § 3.216 (b):Obtain from Curve B. Use distribution of

Figure 3-8.VERTICAL TAIL S URFACES

(3) Condition § 3.219 (a):Obtain as function of W/S and surface

deflection in same manner as outlined in (1)above, use distribution of Figure 3-8;(4) Condition § 3.219 (b):

Obtain from Curve C, use distribution ofFigure 3-7;(5) Condition § 3.219 (c):

Obtain from Curve A, use distribution ofFigure 3-9. (Note that condition § 3.220 generallywill be more critical than this condition.)

AILERONS

(6) In lieu of conditions § 3.222 (b):

Obtain from Curve B, acting in both up anddown directions. Use distribution of Figure 3-10.where:

nm = positive limit maneuvering load factorused in the design of the airplane.

V = initial speed in miles per hour.

FIG. 3-3(a)—LIMIT AVERAGE MANEUVERING CONTROL SURFACE LOADINGS

(d) The total tail load for the conditionsspecified in (c) shall be the sum of: (1) Thebalancing tail load corresponding with thecondition at speed V and the specified value ofthe normal load factor n, plus (2) the maneuveringload increment due to the specified value of theangular acceleration.

NOTE: The maneuvering load increment ofFigure 3-4 and the distributions of Figure 3-8 (fordownloads) and Figure 3-9 (for uploads) may beused. These distributions apply to the total tailload.

§ 3.217 Gust loads. The horizontal tailsurfaces shall be designed for loads occurring inthe following conditions:

(a) Positive and negative gusts of 30 feetper second nominal intensity at speed Vc,corresponding to flight condition § 3.187 (a) withflaps retracted.

NOTE: The average loadings of Figures 3-5(a) and 3-5 (b) and the distribution of Figure 3-9may be used for the total tail loading in thiscondition.

(b) Positive and negative gusts of 15 feetper second nominal intensity at speed Vt,corresponding to flight condition § 3.190 (b) withflaps extended. In determining the total load onthe horizontal tail for these conditions, the initialbalancing tail loads shall first be determined forsteady unaccelerated flight at the pertinentdesign speeds Vc and Vt. The incremental tailload resulting from the gust shall then be addedto the initial balancing tail load to obtain the totaltail load.

Note: The incremental tail load due to thegust may be computed by the following formula:

where:

= the limit gust load increment on the tailin pounds;

K = gust coefficient K in § 3.188,

U = nominal gust intensity in feet per second,

V = airplane speed in miles per hour,

St = tail surface area in square feet,

at = slope of lift curve of tail surface,

CL per degree, corrected for aspect ratio,

aw = slope of lift curve of wing, CL per degree,

Rw = aspect ratio of the wing.

§ 3.218 Unsymmetrical loads. Themaximum horizontal tail surface loading (load perunit area), as determined by the precedingsections, shall be applied to the horizontalsurfaces on one side of the plane of symmetryand the following percentage of that loading shallbe applied on the opposite side:

% = 100-10 (n-1) where:

n is the specified positive maneuvering loadfactor.

In any case the above value shall not begreater than 80 percent.

VERTICAL TAIL S URFACES

§ 3.219 Maneuvering loads. At all speedsup to Vp:

(a) With the airplane in unaccelerated flightat zero yaw, a sudden displacement of the ruddercontrol to the maximum deflection as limited bythe control stops or pilot effort, whichever iscritical, shall be assumed.


(b) The airplane shall be assumed to beyawned to a sideslip angle of 15 degrees whilethe rudder control is maintained at full deflection(except as limited by pilot effort) in the directiontending to increase the sideslip.


(c) The airplane shall be assumed to beyawed to a sideslip angle of 15 degrees while therudder control is maintained in the neutralposition (except as limited by pilot effort). Theassumed sideslip angles may be reduced if isshown that the value chosen for a particularspeed cannot be exceeded in the cases of steadyslips, uncoordinated rolls from a steep bank, andsudden failure of the critical engine with delayedcorrective action.


§ 3.220 Gust loads. (a) The airplane shallbe assumed to encounter a gust of 30 feet per

second nominal intensity, normal to the plane ofsymmetry while in unaccelerated flight at speedVc.

(b) The gust loading shall be computed bythe following formula:

w = KUVm575

where:

w = average limit unit pressure in pounds persquare foot,

K = except that K shall not beless than 1.0. A value of K obtained by rationaldetermination may be used.

U = nominal gust intensity in feet per second,

V = airplane speed in miles per hour,

m = slope of lift curve of vertical surface, CL

per radian, corrected for aspect ratio,

W = design weight in pounds,

Sv = vertical surface area in square feet.

(c) This loading applies only to that portionof the vertical surfaces having a well-definedleading edge.


§ 3.221 Outboard fins. When outboard finsare carried on the horizontal tail surface, the tailsurfaces shall be designed for the maximumhorizontal surface load in combination with thecorresponding loads induced on the verticalsurfaces by end plate effects. Such inducedeffects need not be combined with other verticalsurface loads. When outboard fins extend aboveand below the horizontal surface, the criticalvertical surface loading (load per unit area) asdetermined by §§ 3.219 and 3.220 shall beapplied:

(a) To the portion of the vertical surfacesabove the horizontal surface, and 80 percent ofthat loading applied to the portion below thehorizontal surface,

(b) To the portion of the vertical surfacesbelow the horizontal surface, and 80 percent ofthat loading applied to the portion above thehorizontal surface.

AILERONS, WING FLAPS, TABS, ETC.

§ 3.222 Ailerons. (a) In the symmetricalflight conditions (see §§ 3.183-3.189), the aileronsshall be designed for all loads to which they aresubjected while in the neutral position.

(b) In unsymmetrical flight conditions (see§ 3.191 (a)), the ailerons shall be designed for theloads resulting from the following deflectionsexcept as limited by pilot effort:

(1) At speed Vp it shall be assumed thatthere occurs a sudden maximum displacement ofthe aileron control. (Suitable allowance may bemade for control system deflections.)

(2) When Vc is greater than Vp, the ailerondeflection at Vc shall be that required to producea rate of roll not less than that obtained incondition (1).

(3) At speed Vd the aileron deflection shallbe that required to produce a rate of roll not lessthan one-third of that which would be obtained atthe speed and aileron deflection specified incondition (1).

Note: For conventional ailerons, thedeflections for conditions (2) and (3) may becomputed from:

δ2= δ1; and δ3= δ1;

where:

δ1 = total aileron deflection (sum of bothaileron deflections) in condition (1).

δ2 = total aileron deflection in condition (2).

δ3 = total deflection in condition (3). In theequation for δ3 the 0.5 factor is used instead of0.33 to allow for wing torsional flexibility.

(c) The critical loading on the aileronsshould occur in condition (2) if Vd is less than 2Vc

and the wing meets the torsional stiffness criteria.The normal force coefficient CN for the aileronsmay be taken as 0.04δ, where δ is the deflection ofthe individual aileron in degrees. The criticalcondition for wing torsional loads will dependupon the basic airfoil moment coefficient as wellas the speed, and may be determined as follows:

T3 δ 31)Vd

2

T2 δ 21)Vc

2

where:

T3/T2 is the ratio of wing torsion in condition(b) (3) to that in condition (b) (2).

δ21 and δ31 are the down deflections of theindividual aileron in conditions (b) (2) and (3)respectively.

(d) When T3/T2 is greater than 1.0 condition(b) (3) is critical; when T3/T2 is less than 1.0condition (b) (2) is critical.

(e) In lieu of the above rational conditionsthe average loading of Figure 3-3 and thedistribution of Figure 3-10 may be used.

§ 3.223 Wing flaps. Wing flaps, theiroperating mechanism, and supporting structureshall be designed for critical loads occurring inthe flap-extended flight conditions (see § 3.190)with the flaps extended to any position from fullyretracted to fully extended; except that when anautomatic flap load limiting device is employedthese parts may be designed for criticalcombinations of air speed and flap positionpermitted by the device. (Also see §§ 3.338 and3.339.) The effects of propeller slipstreamcorresponding to take-off power shall be takeninto account at an airplane speed of not less than1.4 Vs where Vs is the computed stalling speedwith flaps fully retracted at the design weight.For investigation of the slipstream condition, theairplane load factor may be assumed to be 1.0.

§ 3.224 Tabs. Control surface tabs shall bedesigned for the most severe combination of airspeed and tab deflection likely to be obtainedwithin the limit V-n diagram (Fig. 3-1) for anyusable loading condition of the airplane.

§ 3.225 Special devices. The loading forspecial devices employing aerodynamic surfaces,such as slots and spoilers, shall be based on testdata.

CONTROL S YSTEM LOADS

§ 3.231 Primary flight controls andsystems. (a) Flight control systems andsupporting structure shall be designed for loadscorresponding to 125 percent of the computedhinge moments of the movable control surface inthe conditions prescribed in §§ 3.211 to 3.225,subject to the following maxima and minima:

(1) The system limit loads need not exceedthose which can be produced by the pilot andautomatic devices operating the controls.

(2) The loads shall in any case be sufficientto provide a rugged system for service use,including consideration of jamming, ground

gusts, taxiing tail to wind, control inertia, andfriction.

(b) Acceptable maximum and minimum pilotloads for elevator, aileron, and rudder controlsare shown in Figure 3-11. These pilot loads shallbe assumed to act at the appropriate control gripsor pads in a manner simulating flight conditionsand to be reacted at the attachments of thecontrol system to the control surface horn.

§ 3.232 Dual controls. When dual controlsare provided, the systems shall be designed forthe pilots operating in opposition, usingindividual pilot loads equal to 75 percent of thoseobtained in accordance with § 3.231, except thatthe individual pilot loads shall not be less thanthe minimum loads specified in Figure 3-11.

§ 3.233 Ground gust conditions. (a) Thefollowing ground gust conditions shall beinvestigated in cases where a deviation from thespecific values for minimum control forces listedin Figure 3-11 is applicable. The followingconditions are intended to simulate the loadingson control surfaces due to ground gusts andwhen taxiing with the wind.

(b) The limit hinge moment H shall beobtained from the following formula:

H = KcSq

where:

H = limit hinge moment (foot-pounds).

c = mean chord of the control surface aft ofthe hinge line (feet).

S = area of control surface aft of the hinge line(square feet).

q = dynamic pressure (pounds per squarefoot) to be based on a design speed not less than

10√W/S+10 miles per hour, except that the designspeed need not exceed 60 miles per hour.

K = factor as specified below:

Surface K(a) Aileron-----------------------------------Control column locked or lashed in mid-position.

+ 0.75

(b) Aileron-----------------------------------Ailerons at full throw; + moment on oneaileron, - moment on the other.

±0.50

(c) (d) Elevator-----------------------------Elevator (c) full up (-), and (d) full down

±0.75

(+).(e) (f) Rudder-------------------------------Rudder (e) in neutral, and (f) at full throw.

±0.75

(c) As used in paragraph (b) in connectionwith ailerons and elevators, a positive value of Kindicates a moment tending to depress thesurface while a negative value of K indicates amoment tending to raise the surface.

§ 3.234 Secondary controls and systems.Secondary controls, such as wheel brakes,spoilers, and tab controls, shall be designed forthe loads based on the maximum which a pilot islikely to apply to the control in question.

GROUND LOADS

§ 3.241 Ground loads. The loads specifiedin the following conditions shall be considered asthe external loads and inertia forces which wouldoccur in an airplane structure if it were acting as arigid body. In each of the ground loadconditions specified the external reactions shallbe placed in equilibrium with the linear andangular inertia forces in a rational or conservativemanner.

§ 3.242 Design weight. The design weightused in the landing conditions shall not be lessthan the maximum weight for which certificationis desired: Provided, however, That formultiengine airplanes meeting the one-engine-inoperative climb requirement of § 3.85 (b), theairplane may be designed for a design landingweight which is less than the maximum designweight, if compliance is shown with the followingsections of Part 4b in lieu of the correspondingrequirements of this part: the ground loadrequirements of § 4b.241, and shock absorptionrequirements of § 4b.371 and its related sections,the wheel and tire requirements of §§ 4b.391 and4b.392, and the fuel jettisoning systemrequirements of § 4b.536.

§ 3.243 Load factor for landing conditions.In the following landing conditions the limitvertical inertia load factor at the center of gravityof the airplane shall be chosen by the designerbut shall not be less than the value which wouldbe obtained when landing the airplane with adescent velocity, in feet per second, equal to thefollowing value:

V = 4.4 (W/S)¼

except that the descent velocity need not exceed10 feet per second and shall not be less than 7

feet per second. Wing lift not exceeding two-thirds of the weight of the airplane may beassumed to exist throughout the landing impactand may be assumed to act through the airplanecenter of gravity. When such wing lift isassumed, the ground reaction load factor may betaken equal to the inertia load factor minus theratio of the assumed wing lift to the airplaneweight. (See § 3.354 for requirements concerningthe energy absorption tests which determine thelimit load factor corresponding to the requiredlimit descent velocities.) In no case, however,shall the inertia load factor used for designpurposes be less than 2.67, nor shall the limitground reaction load factor be less than 2.0,unless it is demonstrated that lower values oflimit load factor will not be exceeded in taxiing theairplane over terrain having the maximum degreeof roughness to be expected under intendedservice use at all speeds up to take-off speed.

LANDING CASES AND ATTITUDES

§ 3.244 Landing cases and attitudes. Forconventional arrangements of main and nose, ormain and tail wheels, the airplane shall beassumed to contact the ground at the specifiedlimit vertical velocity in the attitudes described in§§ 3.245-3.247. (See Figs. 3-12 (a) and 3-12 (b) foracceptable landing conditions which areconsidered to conform with §§ 3.245-3.247.)

§ 3.245 Level landing—(a) Tail wheel type.Normal level flight attitude.

(b) Nose wheel type. Two cases shall beconsidered:

(1) Nose and main wheels contacting theground simultaneously,

(2) Main wheels contacting the ground,nose wheel just clear of the ground. (The

angular attitude may be assumed the same asin subparagraph (1) of this paragraph forpurposes of analysis.)

(c) Drag components. In this condition,drag components simulating the forces requiredto accelerate the tires and wheels up to thelanding speed shall be properly combined with

the corresponding instantaneous vertical groundreactions. The wheel spin-up drag loads may bebased on vertical ground reactions, assumingwing lift and a tire-sliding coefficient of friction of0.8, but in any case the drag loads shall not beless than 25 percent of the maximum verticalground reactions neglecting wing lift.

LIMIT PILOT LOADS

ControlMaximum loads for design weightW equal to or less than 5,000 lbs.1 Minimum loads.2

Aileron: Stick------------------------------- Wheel3----------------------------

67 pounds----------------------------53 D in-pounds4--------------------

40 pounds.40 D in-pounds.

Elevator: Stick------------------------------- Wheel-----------------------------

167 pounds--------------------------200 pounds--------------------------

100 pounds.100 pounds.

Rudder------------------------------- 200 pounds-------------------------- 130 pounds.1For design weight W greater than 5,000 pounds the above specified maximum values shall be increased

linearly with weight to 1.5 times the specified values at a design weight of 25,000 pounds.2If the design of any individual set of control systems or surfaces is such as to make these specified

minimum loads inapplicable, values corresponding to the pertinent binge moments obtained according to §3.233 may be used instead, except that in any case values less than 0.6 of the specified minimum loads shallnot be employed.

3The critical portions of the aileron control system shall also be designed for a single tangential forcehaving a limit value equal to 1.25 times the couple force determined from the above criteria.

4D = wheel diameter.FIG. 3-11 ——PILOT CONTROL FORCE LIMITS

§ 3.246 Tail down—(a) Tail wheel type.Main and tail wheels contacting groundsimultaneously.

(b) Nose wheel type. Stalling attitude or themaximum angle permitting clearance of theground by all parts of the airplane, whichever isthe lesser.

(c) Vertical ground reactions. In thiscondition, it shall be assumed that the groundreactions are vertical, the wheels having beenbrought up to speed before the maximum verticalload is attained.

§ 3.247 One-wheel landing. One side ofthe main gear shall contact the ground with theairplane in the level attitude. The groundreactions shall be the same as those obtained onthe one side in the level attitude. (See § 3.245.)

GROUND ROLL CONDITIONS

§ 3.248 Braked roll. The limit vertical loadfactor shall be 1.33. The attitude and groundcontacts shall be those described for level

landings in § 3.245, with the shock absorbers andtires deflected to their static positions. A dragreaction equal to the vertical reaction at the wheelmultiplied by a coefficient of friction of 0.8 shallbe applied at the ground contact point of eachwheel having brakes, except that the dragreaction need not exceed the maximum valuebased on limiting brake torque.

§ 3.249 Side load. Level attitude with mainwheels only contacting the ground, with theshock absorbers and tires deflected to their staticpositions. The limit vertical load factor shall be1.33 with the vertical ground reaction dividedequally between main wheels. The limit sideinertia factor shall be 0.83 with the side groundreaction divided between main wheels as follows:

0.5 W acting inboard on one side.

0.33 W acting outboard on the other side.

TAIL WHEELS

§ 3.250 Supplementary conditions for tail

wheels. The conditions in §§ 3.251 and 3.252apply to tail wheels and affected supportingstructure.

§ 3.251 Obstruction load. The limit groundreaction obtained in the tail down landingcondition shall be assumed to act up and aftthrough the axle at 45 degrees. The shockabsorber and tire may be assumed deflected totheir static positions.

§ 3.252 Side load. A limit vertical groundreaction equal to the static load on the tail wheel,in combination with a side component of equalmagnitude. When a swivel is provided, the tailwheel shall be assumed swiveled 90 degrees tothe airplane longitudinal axis, the resultantground load passing through the axle. When alock steering device or shimmy damper isprovided, the tail wheel shall also be assumed inthe trailing position with the side load acting atthe ground contact point. The shock absorberand tire shall be assumed deflected to their staticpositions.

NOSE WHEELS

§ 3.253 Supplementary conditions for nosewheels. The conditions set forth in §§ 3.254-3.256 apply to nose wheels and affectedsupporting structure. The shock absorbers andtires shall be assumed deflected to their staticpositions.

§ 3.254 Aft load. Limit force components ataxle:

Vertical, 2.25 times static load on wheel, Drag,0.8 times vertical load.

§ 3.255 Forward load. Limit forcecomponents at axle:

Vertical, 2.25 times static load on wheel,Forward, 0.4 times vertical load.

§ 3.256 Side load. Limit force componentsat ground contact:

Vertical, 2.25 times static load on wheel, Side,0.7 times vertical load.

S KIPLANES

§ 3.257 Supplementary conditions forskiplanes. The airplane shall be assumed restingon the ground with one main ski frozen in thesnow and the other main ski and the tail ski freeto slide. A limit side force equal to P/3 shall beapplied at the most convenient point near the tailassembly, where P is the static ground reactionon the tail ski. For this condition the factor ofsafety shall be assumed equal to 1.0.

WATER LOADS

§ 3.265 General. The requirements setforth in §§ 3.266-3.282 shall apply to the entireairplane, but have particular reference to hullstructure, wing, nacelles, and float supportingstructure.

DESIGN WEIGHT

§ 3.266 Design weight. The design weightused in the water landing conditions shall not beless than the maximum weight for whichcertification is desired for any operation.

BOAT S EAPLANES

§ 3.267 Local bottom pressures—(a)Maximum local pressure. The maximum value ofthe limit local pressure shall be determined fromthe following equation:

where

P = pressure in pound per square inch.

Vs0 = stalling speed, flaps down, power off, inmiles per hour (to be calculated on the basis ofwind tunnel data or flight tests on previousairplanes).

W = design weight.

(b) Variation in local pressure. The localpressures to be applied to the hull bottom shallvary in accordance with Figure 3-13. No variationfrom keel to chine (beamwise) shall be assumed,except when the chine flare indicates theadvisability of higher pressures at the chine.

Tail wheel type Nose wheel type

ConditionLevel

landingTail-down

landingLevel landingwith inclined

reactions

Level landing withnose wheel just clear

of ground

Tail-downlanding

Reference section----- §3.245(a) §3.246(a) § 3.245 (b) (1) § 3.245 (b) (2) § 3.246 (b) (c)Vertical component at c.g-----------------------

nW nW nW nW nW

Fore and aft component at c.g.-------------------

KnW 0 KnW KnW 0

Lateral component in either direction at c.g.

0 0 0 0 0

Shock absorber extension (hydraulic shock absorber)--------- Note (2) Note (2) Note (2) Note (2) Note (2)Shock absorber deflection (rubber or spring shock absorber)---------------- 100% 100% 100% 100% 100%Tire deflection------------ Static Static Static Static StaticMain wheel loads (both wheels)--------------{Vr

{Dr

nWKvr

nW b/d0

nW bf/df

Kvr

nWKvr

nW0

Tail (nose) wheel loads -----------------------{Vf

{Df

00

nW a/d0

nW af/df

Kvf

00

00

Notes---------------------- (1) and (3) ---------------

(1) (1) and (3) (3)

Note (1).—K may be determined as follows: K=0.25 for W=3,000 pounds or less; K=0.33 for W=6,000 pounds or greaterwith linear variation of K between these weights.

Note (2).—For the purpose of design, the maximum load factor shall be assumed to occur throughout the shock absorberstroke from 25 percent deflection to 100 percent deflection unless demonstrated otherwise, and the load factor shall be usedwith whatever shock absorber extension is most critical for each element of the landing gear.

Note (3).—Unbalanced moments shall be balanced by a rational or conservative method.FIG. 3-12(a)——BASIC LANDING CONDITIONS

(c) Application of local pressure. The localpressures determined in paragraphs (a) and (b) ofthis section shall be applied over a local area insuch a manner as to cause the maximum localloads in the hull bottom structure.

§ 3.268 Distributed bottom pressures. (a)For the purpose of designing frames, keels, andchine structure, the limit pressures obtained from§ 3.267 and Figure 3-13 shall be reduced to one-half the local values and simultaneously appliedover the entire hull bottom. The loads soobtained shall be carried into the side-wallstructure of the hull proper, but need not betransmitted in a fore-and-aft direction as shearand bending loads.

(b) Unsymmetrical loading. Each floormember or frame shall be designed for a load onone side of the hull center line equal to the mostcritical symmetrical loading, combined with a loadon the other side of the hull center line equal toone-half of the most critical symmetrical loading.

§ 3.269 Step loading condition—(a)Application of load. The resultant water loadshall be applied vertically in the plane ofsymmetry so as to pass through the center ofgravity of the airplane.

(b) Acceleration. The limit accelerationshall be 4.33.

(c) Hull shear and bending loads. The hullshear and bending loads shall be computed fromthe inertia loads produced by the vertical waterload. To avoid excessive local shear loads andbending moments near the point of water loadapplication, the water load may be distributedover the hull bottom, using pressures not lessthan those specified in § 3.268.

§ 3.270 Bow loading condition—(a)Application of load. The resultant water loadshall be applied in the plane of symmetry at apoint one-tenth of the distance from the bow tothe step and shall be directed upward andrearward at an angle of 30 degrees from thevertical.

(b) Magnitude of load. The magnitude ofthe limit resultant water load shall be determinedfrom the following equation:

where:

Pb = the load in pounds

Ns = the step landing load factor.

We = an effective weight which is assumedequal to one-half the design weight of theairplane.

(c) Hull shear and bending loads. The hullshear and bending loads shall be determined byproper consideration of the inertia loads whichresist the linear and angular accelerationsinvolved. To avoid excessive local shear loads,the water reaction may be distributed over thehull bottom, using pressures not less than thosespecified in § 3.268.

§ 3.271 Stern loading condition—(a)Application of load. The resultant water loadshall be applied vertically in the plane ofsymmetry and shall be distributed over the hullbottom from the second step forward with anintensity equal to the pressures specified in§§ 3.267-3.272.

(b) Magnitude of load. The limit resultantload shall equal three-fourths of the maximumdesign weight of the airplane.

(c) Hull shear and bending loads. The hullshear and bending loads shall be determined byassuming the hull structure to be supported atthe wing attachment fittings and neglectinginternal inertia loads. This condition need not beapplied to the fittings or to the portion of the hullahead of the rear attachment fittings.

§ 3.272 Side loading condition—(a)Application of load. The resultant water loadshall be applied in a vertical plane through thecenter of gravity. The vertical component shallbe assumed to act in the plane of symmetry andhorizontal component at a point halfway betweenthe bottom of the keel and the load water line atdesign weight (at rest).

(b) Magnitude of load. The limit verticalcomponent of acceleration shall be 3.25 and theside component shall be equal to 15 percent ofthe vertical component.

(c) Hull shear and bending loads. The hullshear and bending loads shall be determined byproper consideration of the inertia loads or byintroducing couples at the wing attachmentpoints. To avoid excessive local shear loads, thewater reaction may be distributed over the hullbottom, using pressures not less than those

specified by § 3.268.

FLOAT S EAPLANES

§ 3.273 Landing with inclined reactions.(a) The vertical component of the limit loadfactor shall be 4.2 except that it need not exceed avalue given by the following formula:

n = 3.0 + 0.133 W/S

(b) The propeller axis (or equivalentreference line) shall be assumed to be horizontaland the resultant water reaction to be acting inthe plane of symmetry and passing through thecenter of gravity of the airplane, but inclined sothat its horizontal component is equal to one-fourth of its vertical component. Inertia forcesshall be assumed to act in a direction parallel tothe water reaction.

(c) Factors of safety. For the design of floatattachment members, including the membersnecessary to complete a rigid brace truss throughthe fuselage, the factor of safety shall be 1.85.For the remaining structural members, the factorof safety shall be 1.5.

§ 3.275 Landing with vertical reactions.(a) The limit load factor shall be 4.33 actingvertically, except that it need not exceed a valuegiven by the following formula:

n = 3.0 + 0.133 W/S

(b) The propeller axis (or equivalentreference line) shall be assumed to be horizontal,and the resultant water reaction to be vertical andpassing through the center of gravity of theairplane.

(c) Factors of safety. The factors of safetyshall be the same as those specified in §3.273(c).

§ 3.277 Landing with side load. Thevertical component of the limit load factor shallbe 4.0. The propeller axis (or equivalent referenceline) shall be assumed to be horizontal and theresultant water reaction shall be assumed to be inthe vertical plane which passes through thecenter of gravity of the airplane and isperpendicular to the propeller axis. The verticalload shall be applied through the keel or keels ofthe float or floats and evenly divided between thefloats when twin floats are used. A side loadequal to one-fourth of the vertical load shall beapplied along a line approximately halfwaybetween the bottom of the keel and the level of

the water line at rest. When twin floats are used,the entire side load specified shall be applied tothe float on the side from which the waterreaction originates.

§ 3.278 Supplementary load conditions.Each main float of a float seaplane shall becapable of carrying the following loads whensupported at the attachment fittings as installedon the airplane:

(a) A limit load, acting upward, applied atthe bow end of float and of magnitude equal tothat portion of the airplane weight normallysupported by the particular float.

(b) A limit load, acting upward, applied atthe stern of magnitude equal to 0.8 times thatportion of the airplane weight normally supportedby the particular float.

(c) A limit load, acting upward, applied atthe step and of magnitude equal to 1.5 times thatportion of the airplane weight normally supportedby the particular float.

§ 3.279 Bottom loads. (a) Main seaplanefloat bottoms shall be designed to withstand thefollowing local pressures:

(1) A limit pressure of at least 10 poundsper square inch over that portion of the bottomlying between the first step and a section 25percent of the distance from the step to the bow.

(2) A limit pressure of at least 5 pounds persquare inch over that portion of the bottom lyingbetween the section 25 percent of the distancefrom the strip to the bow and a section 75 percentof the distance from the step to the bow.

(3) A limit pressure of at least 3 pounds persquare inch over that portion of the bottom aft ofthe step (aft of main step if more than one step isused).

(b) The local pressures determined inparagraph (a) (1), (2) and (3) of this section shallbe applied over local areas in such a manner as tocause the maximum loads in local structure suchas bottom plating and stringers.

(c) For the purpose of designing frames,keels, and chine structure, distributed bottompressures equal to one-half of the local valuesspecified above shall be applied over the entirespecified bottom areas.

WING-TIP FLOAT AND S EA WING LOADS

§ 3.280 Wing-tip float loads. Wing-tipfloats and their attachment, including the wingstructure, shall be analyzed for each of thefollowing conditions:

(a) A limit load acting vertically up at thecompletely submerged center of buoyancy andequal to 3 times the completely submergeddisplacement.

(b) A limit load inclined upward at 45degrees to the rear and acting through thecompletely submerged center of buoyancy andequal to 3 times the completely submergeddisplacement.

(c) A limit load acting parallel to the watersurface (laterally) applied at the center of area ofthe side view and equal to 1.5 times thecompletely submerged displacement.

§ 3.281 Wing structure. The primary wingstructure shall incorporate sufficient extrastrength to insure that failure of wing-tip floatattachment members occurs before the wingstructure is damaged.

§ 3.282 Sea wing loads. Sea wing designloads shall be based on suitable test data.

S UBPART D——DESIGN AND CONSTRUCTION

GENERAL

§ 3.291 General. The suitability of allquestionable design details or parts having animportant bearing on safety in operation shall beestablished by tests.

§ 3.292 Materials and workmanship. Thesuitability and durability of all materials used inthe airplane structure shall be established on thebasis of experience or tests. All materials used inthe airplane structure shall conform to approvedspecifications which will insure their having thestrength and other properties assumed in thedesign data. All workmanship shall be of a highstandard.

§ 3.293 Fabrication methods. Themethods of fabrication employed in constructingthe airplane structure shall be such as to produceconsistently sound structure. When afabrication process such as gluing, spot welding,or heat-treating requires close control to attainthis objective, the process shall be performed inaccordance with an approved processspecification.

§ 3.294 Standard fastenings. All bolts,

pins, screws, and rivets used in the structureshall be of an approved type. The use of anapproved locking device or method is requiredfor all such bolts, pins, and screws. Self-lockingnuts shall not be used on bolts subject torotation during the operation of the airplane.

§ 3.295 Protection. All members of thestructure shall be suitably protected againstdeterioration or loss of strength in service due toweathering, corrosion, abrasion, or other causes.In seaplanes, special precaution shall be takenagainst corrosion from salt water, particularlywhere parts made from different metals are inclose proximity. Adequate provisions forventilation and drainage of all parts of thestructure shall be made.

§ 3.296 Inspection provisions. Adequatemeans shall be provided to permit the closeexamination of such parts of the airplane asrequire periodic inspection, adjustments forproper alignment and functioning, and lubricationof moving parts.

S TRUCTURAL PARTS

§ 3.301 Material strength properties anddesign values. Material strength properties shallbe based on a sufficient number of tests ofmaterial conforming to specifications to establishdesign values on a statistical basis. The designvalues shall be so chosen that the probability ofany structure being understrength because ofmaterial variations is extremely remote. Valuescontained in ANC-5 and ANC-18 shall be usedunless shown to be inapplicable in a particularcase.

Note: ANC-5, "Strength of Aircraft Elements"and ANC-18, "Design of Wood AircraftStructures" are published by the Army-Navy-Civil Committee on Aircraft Design Criteria andmay be obtained from the Government PrintingOffice, Washington 25, D.C.

§ 3.302 Special factors. Where there maybe uncertainty concerning the actual strength ofparticular parts of the structure or where thestrength is likely to deteriorate in service prior tonormal replacement, increased factors of safetyshall be provided to insure that the reliability ofsuch parts is not less than the rest of thestructure as specified in §§ 3.303-3.306.

§ 3.303 Variability factor. For parts whosestrength is subject to appreciable variability due

to uncertainties in manufacturing processes andinspection methods, the factor of safety shall beincreased sufficiently to make the probability ofany part being under-strength from this causeextremely remote. Minimum variability factors(only the highest pertinent variability factor needbe considered) are set forth in §§ 3.304-3.306.

§ 3.304 Castings. (a) Where visualinspection only is to be employed, the variabilityfactor shall be 2.0.

(b) The variability factor may be reduced to1.25 for ultimate loads and 1.15 for limit loadswhen at least three sample castings are tested toshow compliance with these factors, and allsample and production castings are visually andradiographically inspected in accordance with anapproved inspection specification.

(c) Other inspection procedures andvariability factors may be used if foundsatisfactory by the Administrator.

§ 3.305 Bearing factors. (a) The factor ofsafety in bearing at bolted or pinned joints shallbe suitably increased to provide for the followingconditions:

(1) Relative motion in operation (controlsurface and system joints are covered in §§3.327-3.347).

(2) Joints with clearance (free fit) subject topounding or vibration.

(b) Bearing factors need not be appliedwhen covered by other special factors.

§ 3.306 Fitting factor. Fittings are deniedas parts such as end terminals used to join onestructural member to another. A multiplyingfactor of safety of at least 1.15 shall be used inthe analysis of all fittings the strength of which isnot proved by limit and ultimate load tests inwhich the actual stress conditions are simulatedin the fitting and the surrounding structure. Thisfactor applies to all portions of the fitting, themeans of attachment, and bearing on themembers joined. In the case of integral fittings,the part shall be treated as a fitting up to thepoint where the section properties becometypical of the member. The fitting factor need notbe applied where a type of joint design based oncomprehensive test data is used. The followingare examples: continuous joints in metal plating,welded joints, and scarf joints in wood, all madein accordance with approved practices.

§ 3.307 Fatigue strength. The structureshall be designed, insofar as practicable, to avoidpoints of stress concentration where variablestresses above the fatigue limit are likely to occurin normal service.

FLUTTER AND VIBRATION

§ 3.311 Flutter and vibration preventionmeasures. Wings, tail, and control surfaces shallbe free from flutter, airfoil divergence, and controlreversal from lack of rigidity, for all conditions ofoperation within the limit V-n envelope, and thefollowing detail requirements shall apply:

(a) Adequate wing torsional rigidity shallbe demonstrated by tests or other methods foundsuitable by the Administrator.

(b) The mass balance of surfaces shall besuch as to preclude flutter.

(c) The natural frequencies of all mainstructural components shall be determined byvibration tests or other methods foundsatisfactory by the Administrator.

WINGS

§ 3.317 Proof of strength. The strength ofstressed-skin wings shall be substantiated byload tests or by combined structural analysis andtests.

§ 3.318 Ribs. (a) The strength of ribs inother than stressed-skin wings shall be provedby test to at least 125 percent of the ultimateloads for the most severe loading conditions,unless a rational load analysis and test procedureis employed and the tests cover the variability ofthe particular type of construction.

(b) The effects of ailerons and high liftdevices shall be properly accounted for. Ribtests shall simulate conditions in the airplanewith respect to torsional rigidity of spars, fixityconditions, lateral support, and attachment tospars.

§ 3.319 External bracing. When wires areused for external lift bracing they shall be doubleunless the design provides for a lift-wire-cutcondition. Rigging loads shall be taken intoaccount in a rational or conservative manner.The end connections of brace wires shall be suchas to minimize restraint against bending orvibration. When brace struts of large finenessratio are used, the aerodynamic forces on such

struts shall be taken into account.

§ 3.320 Covering. Strength tests of fabriccovering shall be required unless approvedgrades of cloth, methods of support, attachment,and finishing are employed. Special tests shall berequired when it appears necessary to accountfor the effects of unusually high design airspeeds, slipstream velocities, or other unusualconditions.

CONTROL S URFACES (FIXED AND MOVABLE)

§ 3.327 Proof of strength. Limit load testsof control surfaces are required. Such tests shallinclude the horn or fitting to which the controlsystem is attached. In structural analyses,rigging loads due to wire bracing shall be takeninto account in a rational or conservative manner.

§ 3.328 Installation. Movable tail surfacesshall be so installed that there is no interferencebetween the surfaces or their bracing when eachis held in its extreme position and all others areoperated through their full angular movement.When an adjustable stabilizer is used, stops shallbe provided which, in the event of failure of theadjusting mechanism, will limit its travel to arange permitting safe flight and landing.

§ 3.329 Hinges. Control surface hinges,excepting ball and roller bearings, shallincorporate a multiplying factor of safety of notless than 6.67 with respect to the ultimate bearingstrength of the softest material used as a bearing.For hinges incorporating ball or roller bearings,the approved rating of the bearing shall not beexceeded. Hinges shall provide sufficientstrength and rigidity for loads parallel to thehinge line.

CONTROL S YSTEMS

§ 3.335 General. All controls shall operatewith sufficient ease, smoothness, andpositiveness to permit the proper performance oftheir function and shall be so arranged andidentified as to provide convenience in operationand prevent the possibility of confusion andsubsequent inadvertent operation. (See § 3.384for cockpit controls.)

§ 3.336 Primary flight controls. (a)Primary flight controls are defined as those usedby the pilot for the immediate control of thepitching, rolling, and yawing of the airplane.

(b) For two-control airplanes the design

shall be such as to minimize the likelihood ofcomplete loss of the lateral directional control inthe event of failure of any connecting ortransmitting element in the control system.

§ 3.337 Trimming controls. Properprecautions shall be taken against the possibilityof inadvertent, improper, or abrupt taboperations. Means shall be provided to indicateto the pilot the direction of control movementrelative to airplane motion and the position of thetrim device with respect of the range ofadjustment. The means used to indicate thedirection of the control movement shall beadjacent to the control, and the means used toindicate the position of the trim device shall beeasily visible to the pilot and so located andoperated as to preclude the possibility ofconfusion. Trimming devices shall be capable ofcontinued normal operation notwithstanding thefailure of any one connecting or transmittingelement in the primary flight control system. Tabcontrols shall be irreversible unless the tab isproperly balanced and possesses no unsafeflutter characteristics. Irreversible tab systemsshall provide adequate rigidity and reliability inthe portion of the system from the tab to theattachment of the irreversible unit to the airplanestructure.

§ 3.338 Wing flap controls. The controlsshall be such that when the flap has been placedin any position upon which compliance with theperformance requirements is based, the flap willnot move from that position except upon furtheradjustment of the control or the automaticoperation of a flap load limiting device. Meansshall be provided to indicate the flap position tothe pilot. If any flap position other than fullyretracted or extended is used to show compliancewith the performance requirements, such meansshall indicate each such position. The rate ofmovement of the flaps in response to theoperation of the pilot’s control, or of anautomatic device shall not be such as to result inunsatisfactory flight or performancecharacteristics under steady or changingconditions of air speed, engine power, andairplane attitude (See § 3.109 (b) and (c).)

§ 3.339 Flap interconnection. (a) Themotion of flaps on opposite sides of the plane ofsymmetry shall be synchronized by a mechanicalinterconnection, unless the airplane isdemonstrated to have safe flight characteristics

while the flaps are retracted on one side andextended on the other.

(b) Where an interconnection is used, in thecase of multinengine airplanes, it shall bedesigned to account for the unsymmetrical loadsresulting from flight with the engines on one sideof the plane of symmetry inoperative and theremaining engines at take-off power. For single-engine airplanes, it may be assumed that 100percent of the critical air load acts on one sideand 70 percent on the other.

§ 3.340 Stops. All control systems shall beprovided with stops which positively limit therange of motion of the control surfaces. Stopsshall be so located in the system that wear,slackness, or take-up adjustments will notappreciably affect the range of surface travel.Stops shall be capable of withstanding the loadscorresponding to the design conditions for thecontrol system.

§ 3.341 Control system locks. When adevice is provided for locking a control surfacewhile the airplane is on the ground or water:

(a) The locking device shall be so installedas to provide unmistakable warning to the pilotwhen it is engaged, and

(b) Means shall be provided to preclude thepossibility of the lock becoming engaged duringflight.

§ 3.342 Proof of strength. Tests shall beconducted to prove compliance with limit loadrequirements. The direction of test loads shall besuch as to produce the most severe loading ofthe control system structure. The tests shallinclude all fittings, pulleys, and brackets used toattach the control system to the primarystructure. Analyses or individual load tests shallbe conducted to demonstrate compliance withthe multiplying factor of safety requirementsspecified for control system joints subjected toangular motion.

§ 3.343 Operation test. An operation testshall be conducted by operating the controlsfrom the pilot compartment with the entire systemso loaded as to correspond to the limit air loadson the surface. In this test there shall be nojamming, excessive friction, or excessivedeflection.

CONTROL S YSTEM DETAILS

§ 3.344 General. All control systems and

operating devices shall be so designed andinstalled as to prevent jamming, chafing, orinterference as a result of inadequate clearancesor from cargo, passengers, or loose objects.Special precautions shall be provided in thecockpit to prevent the entry of foreign objectsinto places where they might jam the controls.Provisions shall be made to prevent the slappingof cables or tubes against parts of the airplane.

§ 3.345 Cable systems. Cables, cablefittings, turnbuckles, splices, and pulleys shall bein accordance with approved specifications.Cables smaller than 1/8-inch diameter shall not beused in primary control systems. The design ofcable systems shall be such that there will not behazardous change in cable tension throughoutthe range of travel under operating conditionsand temperature variations. Pulley types andsizes shall correspond to the cables with whichthey are used, as specified on the pulleyspecification. All pulleys shall be provided withsatisfactory guards which shall be closely fittedto prevent the cables becoming misplaced orfouling, even when slack. The pulleys shall lie inthe plane passing through the cable within suchlimits that the cable does not rub against thepulley flange. Fairleads shall be so installed thatthey are not required to cause a change in cabledirection of more than 3 degrees. Clevis pins(excluding those not subject to load or motion)retained only by cotter pins shall not beemployed in the control system. Turnbucklesshall be attached to parts having angular motionin such a manner as to prevent positively bindingthroughout the range of travel. Provisions forvisual inspection shall be made at all fairleads,pulleys, terminals, and turnbuckles.

§ 3.346 Joints. Control system jointssubject to angular motion in push-pull systems,excepting ball and roller bearing systems, shallincorporate a multiplying factor of safety of notless than 3.33 with respect to the ultimate bearingstrength of the softest material used as a bearing.This factor may be reduced to 2.0 for such jointsin cable control systems. For ball or rollerbearings the approved rating of the bearing shallnot be exceeded.

§ 3.347 Spring devices. The reliability ofany spring devices used in the control systemshall be established by tests simulating serviceconditions, unless it is demonstrated that failureof the spring will not cause flutter or unsafe flight

characteristics.

LANDING GEAR

S HOCK ABSORBERS

§ 3.351 Tests. Shock absorbing elements inmain, nose, and tail wheel units shall besubstantiated by the tests specified in thefollowing section. In addition, the shockabsorbing ability of the landing gear in taxiingmust be demonstrated in the operational tests of§ 3.146.

§ 3.352 Shock absorption tests. (a) It shallbe demonstrated by energy absorption tests thatthe limit load factors selected for design inaccordance with § 3.243 will not be exceeded inlandings with the limit descent velocity specifiedin that section.

(b) In addition, a reserve of energyabsorption shall be demonstrated by a test inwhich the descent velocity is at least 1.2 timesthe limit descent velocity. In this test there shallbe no failure of the shock aborting unit, althoughyielding of the unit will be permitted. Wing liftequal to the weight of the airplane may beassumed for purposes of this test.

§ 3.353 Limit drop tests. (a) If compliancewith the specified limit landing conditions of §3.352 (a) is demonstrated by free drop tests, theseshall be conducted on the complete airplane, oron units consisting of wheel, tire, and shockabsorber in their proper relation, from free dropheights not less than the following:

h (inches) = 3.6 (W/S)0.5

except that the free drop height shall not beless than 9.2 inches and need not be greater than18.7 inches.

(b) In simulating the permissible wing lift infree drop tests, the landing gear unit shall bedropped with an effective mass equal to:

where

We = the effective weight to be used in thedrop test.

h = specified height of drop in inches.

d = deflection under impact of the tire (at theapproved inflation pressure) plus the verticalcomponent of the axle travel relative to the dropmass. The value of d used in the computation of

We shall not exceed the value actually obtained inthe drop tests.

W = WM or main gear units, and shall be equalto the static weight on the particular unit with theairplane in the level attitude (with the nose wheelclear, in the case of nose wheel clear, in the caseof nose wheel type airplanes).

W = WT for tail gear units, and shall be equalto the static weight on the tail unit with theairplane in the tail down attitude.

W = WN for nose wheel units, and shall beequal to the static reaction which will exist at thenose wheel when the mass of the airplane isconcentrated at the center of gravity and exerts aforce of 1.0g downward and 0.33g forward.

L = ratio of assumed wing lift to airplaneweight, not greater than 0.667.

The attitude in which the landing gear unit isdrop tested shall be such as to simulate theairplane landing condition which is critical fromthe standpoint of energy to be absorbed by theparticular unit.

§ 3.354 Limit load factor determination. Indetermining the limit airplane inertia load factor nfrom the free drop test described above, thefollowing formula shall be used:

where

nj = the load factor developed in the drop test,i.e., the acceleration (dv/dt) in g’s recorded in thedrop test, plus 1.0.

The value of n so determined shall not begreater than the limit inertia load factor used inthe landing conditions, § 3.243.

§ 3.355 Reserve energy absorption droptests. If compliance with the reserve energyabsorption condition specified in § 3.352 (b) isdemonstrated by free drop tests, the drop heightshall be not less than 1.44 times the drop heightspecified in § 3.353. In simulating wing lift equaltot he airplane weight, the units shall be droppedwith an effective mass equal to

where the symbols and other details are thesame as in § 3.353.

RETRACTING MECHANISM

§ 3.356 General. The landing gearretracting mechanism and supporting structureshall be designed for the maximum load factors inthe flight conditions when the gear is in theretracted position. It shall also be designed forthe combination of friction, inertia, brake torque,and air loads occurring during retraction at anyair speed up to 1.6Vs1, flaps retracted and anyload factors up to those specified for the flapsextended condition, § 3.190. The landing gearand retracting mechanism, including the wheelwell doors, shall withstand flight loads with thelanding gear extended at any speed up to at least1.6 Vs1 flaps retracted. Positive means shall beprovided for the purpose of maintaining thewheels in the extended position.

§ 3.357 Emergency operation. When otherthan manual power for the operation of thelanding gear is employed, an auxiliary means ofextending the landing gear shall be provided.

§ 3.358 Operation test. Proper functioningof the landing gear retracting mechanism shall bedemonstrated by operation tests.

§ 3.359 Position indicator and warningdevice. When retractable landing wheels areused, means shall be provided for indicating tothe pilot when the wheels are secured in theextreme positions. In addition, landplanes shallbe provided with an aural or equally effectivewarning device which shall functioncontinuously after the throttle is closed until thegear is down and locked.

§ 3.360 Control. See § 3.384.

WHEELS AND TIRES

§ 3.361 Wheels. (a) Main landing gearwheels (i.e., those nearest the airplane center ofgravity) shall be of an approved type.

(b) The rated static load of each main wheelshall not be less than the design weight forground loads (§ 3.242) divided by the number ofmain wheels. Nose wheels shall have been testedfor an ultimate radial load not less than themaximum nose wheel ultimate load obtained inthe ground loads requirements, and forcorresponding side and burst loads.

§ 3.362 Tires. A landing gear wheel may beequipped with any make or type of tire, providedthat the approved tire rating is not exceededunder the following conditions:

(a) Load on main wheel tires equal to theairplane weight divided by the number of wheels,

(b) Load on nose wheel tires (to becompared with the dynamic rating established forsuch tires) equal to the reaction obtained at thenose wheel, assuming the mass of the airplaneconcentrated at the center of gravity and exertinga force of 1.0g downward and 0.31g forward, thereactions being distributed to the nose and mainwheels by the principle of statics with the dragreaction at the ground applied only at thosewheels having brakes. When speciallyconstructed tires are used to support an airplane,the wheels shall be plainly and conspicuouslymarked to that effect. Such marking shall includethe make, size, number of plies, and identificationmarking of the proper tire.

Note: Approved ratings are those assignedby the Tire and Rim Association or by theAdministrator.

BRAKES

§ 3.363 Brakes. Brakes shall be installedwhich are adequate to prevent the airplane fromrolling on a paved runway while applying take-offpower to the critical engine, and of sufficientcapacity to provide adequate speed controlduring taxiing without the use of excessive pedalor hand forces.

S KIS

§ 3.364 Skis. Skis shall be of an approvedtyped. The approved rating of the skis shall notbe less than the maximum weight of the airplaneon which they are installed.

§ 3.365 Installation. (a) When typecertificated skis are installed, the installation shallbe made in accordance with the ski or airplanemanufacturer’s recommendations which shallhave been approved by the Administrator.When other than type certificated skis areinstalled, data shall be submitted to theAdministrator showing a dimensional drawing ofthe proposed method of attaching the skis, thesizes and material of the restraining members andattachment fittings.

(b) In addition to such shock cord(s) asmay be provided, front and rear check cablesshall be used on skis not equipped with specialstabilizing devices.

§ 3.366 Tests. (a) If the airplane is of amodel not previously approved with the specific

ski installation, it shall satisfactorily pass aground inspection of the installation,demonstrate satisfactory landing and taxiingcharacteristics, and comply with such flight testsas are found necessary to indicate that theairplane’s flight characteristics are satisfactorywith the skis installed.

(b) If the airplane is of a model previouslyapproved with the specific ski installation, it needpass satisfactorily only a ground inspection ofthe installation.

HULLS AND FLOATS

§ 3.371 Buoyancy (main seaplane floats).(a) Main seaplane floats shall have a buoyancyin excess of that required to support the maximumweight of the airplane in fresh water as follows:

(1) 80 percent in the case of single floats.

(2) 90 percent in the case of double floats.

(b) Main seaplane floats for use onairplanes of 2,500 pounds or more maximumweight shall contain at least 5 watertightcompartments of approximately equal volume.Main seaplane floats for use on airplanes of lessthan 2,500 pounds maximum weight shall containat least four such compartments.

§ 3.372 Buoyancy (boat seaplanes). Thehulls of boat seaplanes and amphibians shall bedivided into watertight compartments inaccordance with the following requirements:

(a) In airplanes of 5,000 pounds or moremaximum weight, the compartments shall be soarranged that, with any two adjacentcompartments flooded, the hull and auxiliaryfloats (and tires, if used) will retain sufficientbuoyancy to support the maximum weight of theairplane in fresh water.

(b) In airplanes of 1,500 to 5,000 poundsmaximum weight, the compartments shall be soarranged that, with any one compartmentflooded, the hull and auxiliary floats (and tires, ifused) will retain sufficient buoyancy to supportthe maximum weight of the airplane in fresh water.

(c) In airplanes of less than 1,500 poundsmaximum weight, watertight subdivision of thehull is not required.

(d) Bulkheads may have watertight doorsfor the purpose of communication betweencompartments.

§ 3.373 Water stability. Auxiliary floatsshall be so arranged that when completelysubmerged in fresh water, they will provide arighting moment which is at least 1.5 times theupsetting moment caused by the airplane beingtilted. A greater degree of stability may berequired by the Administrator in the case of largeflying boats, depending on the height of thecenter of gravity above the water level, the areaand location of wings and tail surfaces, and otherconsiderations.

FUSELAGE

PILOT COMPARTMENT

§ 3.381 General. (a) The arrangement ofthe pilot compartment and its appurtenancesshall provide a satisfactory degree of safety andassurance that the pilot will be able to perform allhis duties and operate the controls in the correctmanner without unreasonable concentration andfatigue.

(b) The primary flight control units listed onFigure 3-14, excluding cables and control rods,shall be so located with respect to the propellersthat no portion of the pilot or controls lies in theregion between the plane of rotation of anyinboard propeller and the surface generated by aline passing through the center of the propellerhub and making an angle of 5° forward or aft ofthe plane of rotation of the propeller.

§ 3.382 Vision. The pilot compartment shallbe arranged to afford the pilot a sufficientlyextensive, clear, and undistorted view for the safeoperation of the airplane. During flight in amoderate rain condition, the pilot shall have anadequate view of the flight path in normal flightand landing, and have sufficient protection fromthe elements so that his vision is not undulyimpaired. This may be accomplished byproviding an openable window or by a means formaintaining a portion of the windshield in a clearcondition without continuous attention by thepilot. The pilot compartment shall be free of glareand reflections which would interfere with thepilot’s vision. For airplanes intended for nightoperation, the demonstration of these qualitiesshall include night flight tests.

§ 3.383 Pilot windshield and windows. Allglass panes shall be of a nonsplintering safetytype.

§ 3.384 Cockpit controls. (a) All cockpitcontrols shall be so located and, except for those

the function of which is obvious, identified as toprovide convenience in operation includingprovisions to prevent the possibility ofconfusion and consequent inadvertentoperations. (See Fig. 3-14 for required sense ofmotion of cockpit controls.) The controls shallbe so located and arranged that when seated itwill be readily possible for the pilot to obtain fulland unrestricted movement of each controlwithout interference form either his clothing orthe cockpit structure.

(b) Identical power-plant controls for theseveral engines in the case of multiengineairplanes shall be so located as to prevent anymisleading impression as to the engines of whichthey relate.

Control Movement and actuationPrimary: Aileron-------- Right (clockwise) for right wing

down. Elevator--------- Rearward for nose up. Rudder--------- Right pedal forward for nose

right.Power plant: Throttle------------------ Forward to open.

Figure 3-14 Cockpit Controls

§ 3.385 Instruments and markings. See §3.661 relative to instrument arrangement. Theoperational markings, instructions, and placardsrequired for the instruments and controls arespecified in §§ 3.756 to 3.765.

EMERGENCY PROVISIONS

§ 3.386 Protection. The fuselage shall bedesigned to give reasonable assurance that eachoccupant, if he makes proper use of belts orharness for which provisions are made in thedesign, will not suffer serious injury during minorcrash conditions as a result of contact of anyvulnerable part of his body with any penetratingor relatively solid object, although it is acceptedthat parts of the airplane may be damaged.

(a) The ultimate accelerations to whichoccupants are assumed to be subjected shall beas follows:

N, U AUpward---------------- 3.0g 4.5gForward--------------- 9.0g 9.0gSideward--------------- 1.5g 1.5g

(b) For airplanes having retractable landinggear, the fuselage in combination with otherportions of the structure shall be designed to

afford protection of the occupants in a wheels-uplanding with moderate descent velocity.

(c) If the characteristics of an airplane aresuch as to make a turn-over reasonably probable,the fuselage of such an airplane in combinationwith other portions of the structure shall bedesigned to afford protection of the occupants ina complete turn-over.

Note: In § 3.386 (b) and (c), a vertical ultimateacceleration of 3g and a friction coefficient of 0.5at the ground may be assumed.

§ 3.387 Exits. (a) Closed cabins onairplanes carrying more than 5 persons shall beprovided with emergency exits consisting ofmovable windows or panels or of additionalexternal doors which provide a clear andunobstructed opening, the minimum dimensionsof which shall be such that a 19-by-26-inchellipse may be completely inscribed therein. Theexits shall be readily accessible, shall not requireexceptional agility of a person using them, andshall be distributed so as to facilitate egresswithout crowding in all probable attitudesresulting from a crash. The method of openingshall be simple and obvious, and the exits shallbe so arranged and marked as to be readilylocated and operated even in darkness.Reasonable provisions shall be made against thejamming of exits as a result of fuselagedeformation. The proper functioning of exitsshall be demonstrated by tests.

(b) The number of emergency exits requiredis as follows:

(1) Airplanes with a total seating capacityof more than 5 persons, but not in excess of 15,shall be provided with at least one emergency exitor one suitable door in addition to the main doorspecified in § 3.389. This emergency exit, orsecond door, shall be on the opposite side of thecabin from the main door.

(2) Airplanes with a seating capacity ofmore than 15 persons shall be provided withemergency exits or doors in addition to thoserequired in paragraph (b) (1) of this section.There shall be one such additional exit or doorlocated either in the top or side of the cabin forevery additional 7 persons or fraction thereofabove 15, except that not more than four exits,including doors, will be required if thearrangement and dimensions are suitable forquick evacuation of all occupants.

(c) If the pilot compartment is separatedfrom the cabin by a door which is likely to blockthe escape in the event of a minor crash, it shallhave its own exit, but such exit shall not beconsidered as an emergency exit for thepassengers.

(d) In categories U and A exits shall beprovided which will permit all occupants to bailout quickly with parachutes.

§ 3.388 Fire precautions—(a) Cabininteriors. Only materials which are flash-resistant shall be used. In compartments wheresmoking is to be permitted, the materials of thecabin lining, floors, upholstery, and furnishingsshall be flame-resistant. Such compartments shallbe equipped with an adequate number of self-contained ash trays. All other compartmentsshall be placarded against smoking.

(b) Combustion heaters. Gasoline operatedcombustion heater installations shall comply withapplicable parts of the power-plant installationrequirements covering fire hazards andprecautions. All applicable requirementsconcerning fuel tanks, lines, and exhaust systemsshall be considered.

PERSONNEL AND CARGO ACCOMMODATIONS

§ 3.389 Doors. Closed cabins on allairplanes carrying passengers shall be providedwith at least one adequate and easily accessibleexternal door. No passenger door shall be solocated with respect to the propeller discs as toendanger persons using the door.

§ 3.390 Seats and berths—(a) Passengerseats and berths. All seats and berths andsupporting structure shall be designed for apassenger weight of 170 pounds (190 poundswith parachute for the acrobatic and utilitycategories) and the maximum load factorscorresponding to all specified flight and groundload conditions including the emergencyconditions of § 3.386.

(b) Pilot seats. Pilot seats shall bedesigned for the reactions resulting from theapplication of the pilot forces to the primary flightcontrols as specified in § 3.231.

(c) Categories U and A. All seats designedto be occupied in the U and A categories under §3.74 (c) (4) shall be designed to accommodatepassengers wearing parachutes.

§ 3.391 Safety belt or harness provisions.

Provisions shall be made at all seats and berthsfor the installation of belts or harness ofsufficient strength to comply with the emergencyconditions of § 3.386.

§ 3.392 Cargo compartments. Each cargocompartment shall be designed for the placardedmaximum weight of contents and critical loaddistributions at the appropriate maximum loadfactors corresponding to all specified flight andground load conditions. Suitable provisionsshall be made to prevent the contents of cargocompartments form becoming a hazard byshifting. Such provisions shall be adequate toprotect the passengers from injury by thecontents of any cargo compartment when theultimate forward acting accelerating force is 4.5g.

§ 3.393 Ventilation. All passenger andcrew compartments shall be suitably ventilated.Carbon monoxide concentration shall not exceed1 part in 20,000 parts of air.

MISCELLANEOUS

§ 3.401 Leveling marks. Leveling marksshall be provided for leveling the airplane on theground.

S UBPART E——POWER-PLANT INSTALLATIONS;RECIPROCATING ENGINES

GENERAL

§ 3.411 Components. (a) The power plantinstallation shall be considered to include allcomponents of the airplane which are necessaryfor its propulsion. It shall also be considered toinclude all components which affect the controlof the major propulsive units or which affect theircontinued safety of operation.

(b) All components of the power-plantinstallation shall be constructed, arranged, andinstalled in a manner which will assure thecontinued safe operation of the airplane andpower plant. Accessibility shall be provided topermit such inspection and maintenance as isnecessary to assure continued airworthiness.

ENGINES AND PROPELLERS

§ 3.415 Engines. Engines installed incertificated airplanes shall be of a type which hasbeen certificated in accordance with theprovisions of Part 13 of this chapter.

§ 3.416 Propellers. (a) Propellers installedin certificated airplanes shall be of a type whichhas been certificated in accordance with theprovisions of Part 14 of this chapter.

(b) The maximum engine power andpropeller shaft rotational speed permissible foruse in the particular airplane involved shall notexceed the corresponding limits for which thepropeller has been certificated.

§ 3.417 Propeller vibration. In the case ofairplanes equipped with metal propellers, themagnitude of the propeller blade vibrationstresses under all normal conditions of operationshall be determined by actual measurements orby comparison with similar installations for whichsuch measurements have been made. Thevibration stresses thus determined shall notexceed values which have been demonstrated tobe safe for continuous operation. Vibration testsmay be waived and the propeller installationaccepted on the basis of service experience,engine or ground tests which show adequatemargins of safety, or other considerations whichsatisfactorily substantiate its safety in thisrespect. In addition to metal propellers, theAdministrator may require that similarsubstantiation of the vibration characteristics beaccomplished for other types of propellers, withthe exception of conventional fixed-pitch woodpropellers.

§ 3.418 Propeller pitch and speedlimitations. The propeller pitch and speed shallbe limited to values which will assure safeoperation under all normal conditions ofoperation and will assure compliance with theperformance requirements specified in §§ 3.81-3.86.

§ 3.419 Speed limitations for fixed-pitchpropellers, ground adjustable pitch propellers,and automatically varying pitch propellerswhich cannot be controlled in flight, (a) Duringtake-off and initial climb at best rate-of-climbspeed, the propeller, in the case of fixed pitch orground adjustable types, shall restrain the engineto a speed not exceeding its maximum permissibletake-off speed and, in the case of automaticvariable-pitch types, shall limit the maximumgoverned engine revolutions per minute to aspeed not exceeding the maximum permissibletake-off speed. In demonstrating compliancewith this provision the engine shall be operatedat full throttle or the throttle settingcorresponding to the maximum permissible take-off manifold pressure.

(b) During a closed throttle glide at theplacard, "never-exceed speed" (see § 3.739), the

propeller shall not cause the engine to rotate at aspeed in excess of 110 percent of its maximumallowable continuous speed.

§ 3.420 Speed and pitch limitations forcontrollable pitch propellers without constantspeed controls. The stops or other meansincorporated in the propeller mechanism torestrict the pitch range shall limit (a) the lowestpossible blade pitch to a value which will assurecompliance with the provisions of § 3.419 (a), and(b) the highest possible blade pitch to a value notlower than the flattest blade pitch with whichcompliance with the provisions of § 3.419 (b) canbe demonstrated.

§ 3.421 Variable pitch propellers withconstant speed controls. (a) Suitable meansshall be provided at the governor to limit thespeed of the propeller. Such means shall limit themaximum governed engine speed to a value notexceeding its maximum permissible take-offrevolutions per minute.

(b) The low pitch blade stop, or othermeans incorporated in the propeller mechanism torestrict the pitch range, shall limit the speed ofthe engine to a value not exceeding 103 percentof the maximum permissible take-off revolutionsper minute under the following conditions:

(1) Propeller blade set in the lowestpossible pitch and the governor inoperative.

(2) Engine operating at take-off manifoldpressure with the airplane stationary and with nowind.

§ 3.422 Propeller clearance. With theairplane loaded to the maximum weight and mostadverse center of gravity position and thepropeller in the most adverse pitch position,propeller clearances shall not be less than thefollowing, unless smaller clearances are properlysubstantiated for the particular design involved:

(a) Ground clearance. (1) Seven inches(for airplanes equipped with nose wheel typelanding gears) or 9 inches (for airplanes equippedwith tail wheel type landing gears) with thelanding gear statically deflected and the airplanein the level normal take-off, or taxiing attitude,whichever is most critical.

(2) In addition to subparagraph (1) of thisparagraph, there shall be positive clearancebetween the propeller and the ground when, withthe airplane in the level take-off attitude, the

critical tire is completely deflated and thecorresponding landing gear strut is completelybottomed.

(b) Water clearance. A minimum clearanceof 18 inches shall be provided unless compliancewith § 3.147 can be demonstrated with lesserclearance.

(c) Structural clearance. (1) One inchradial clearance between the blade tips and theairplane structure, or whatever additional radialclearance is necessary to preclude harmfulvibration of the propeller or airplane.

(2) One-half inch longitudinal clearancebetween the propeller blades or cuffs andstationary portions of the airplane. Adequatepositive clearance shall be provided betweenother rotating portions of the propeller or spinnerand stationary portions of the airplane.

FUEL S YSTEM

§ 3.429 General. The fuel system shall beconstructed and arranged in a manner to assurethe provision of fuel to each engine at a flow rateand pressure adequate for proper enginefunctioning under all normal conditions ofoperation, including all maneuvers and acrobaticsfor which the airplane is intended.

ARRANGEMENT

§ 3.430 Fuel system arrangement. Fuelsystems shall be so arranged as to permit anyone fuel pump to draw fuel from only one tank ata time. Gravity feed systems shall not supply fuelto any one engine from more than one tank at atime unless the tank air spaces are interconnectedin such a manner as to assure that allinterconnected tanks will feed equally. (See also§ 3.439.)

§ 3.431 Multiengine fuel systemarrangement. The fuel systems of multiengineairplanes shall be arranged to permit operation insuch a manner that the failure of any onecomponent will not result in the loss of the powerof more than one engine. Unless otherprovisions are made in order to comply with thisrequirement, the fuel system shall be arranged topermit supplying fuel to each engine through asystem entirely independent of any portion of thesystem supplying fuel to the other engines.

§ 3.432 Pressure cross feed arrangements.Pressure cross feed lines shall not pass throughportions of the airplane devoted to carrying

personnel or cargo, unless means are provided topermit the flight personnel to shut off the supplyof fuel to these lines, or unless any joints,fittings, or other possible sources of leakageinstalled in such lines are enclosed in a fuel- andfume-proof enclosure which is ventilated anddrained to the exterior of the airplane. Baretubing need not be enclosed but shall beprotected where necessary against possibleinadvertent damage.

OPERATION

§ 3.433 Fuel flow rate. The ability of thefuel system to provide the required fuel flow rateand pressure shall be demonstrated when theairplane is in the attitude which represents themost adverse condition from the standpoint offuel feed and quantity of unusable fuel in thetank. During this test fuel shall be delivered tothe engine at the applicable flow rate (see §§3.434-3.436) and at a pressure not less than theminimum required for proper carburetoroperation. A suitable mock-up of the system, inwhich the most adverse conditions are simulated,may be used for this purpose. The quantity offuel in the tank being tested shall not exceed theamount established as the unusable fuel supplyfor that tank as determined by demonstration ofcompliance with the provisions of § 3.437 (seealso §§ 3.440 and 3.672), plus whatever minimumquantity of fuel it may be necessary to add forthe purpose of conducting the flow test. If a fuelflowmeter is provided, the meter shall be blockedduring the flow test and the fuel shall flowthrough the meter bypass.

§ 3.434 Fuel flow rate for gravity feedsystems. The fuel flow rate for gravity feedsystems (main and reserve supply) shall be 1.2pounds per hour for each take-off horsepower or150 percent of the actual take-off fuelconsumption of the engine, whichever is greater.

§ 3.435 Fuel flow rate for pump systems.The fuel flow rate for pump systems (main andreserve supply) shall be 0.9 pound per hour foreach take-off horsepower or 125 percent of theactual take-off fuel consumption of the engine,whichever is greater. This flow rate shall beapplicable to both the primary engine-drivenpump and the emergency pumps and shall beavailable when the pump is running at the speedat which it would normally be operating duringtake-off. In the case of hand-operated pumps,this speed shall be considered to be not more

than 60 complete cycles (120 single strokes) perminute.

§ 3.436 Fuel flow rate for auxiliary fuelsystems and fuel transfer systems. Theprovisions of § 3.434 or § 3.435, whichever isapplicable, shall also apply to auxiliary andtransfer systems with the exception that therequired fuel flow rate shall be established uponthe basis of maximum continuous power andspeed instead of take-off power and speed. Alesser flow rate shall be acceptable, however, inthe case of a small auxiliary tank feeding into alarge main tank, provided a suitable placard isinstalled to require that the auxiliary tank mustonly be opened to the main tank when apredetermined satisfactory amount of fuel stillremains in the main tank.

§ 3.437 Determination of unusable fuelsupply and fuel system operation on low fuel. (a)The unusable fuel supply for each tank shall beestablished as not less than the quantity at whichthe first evidence of malfunctioning occurs underthe conditions specified in this section. (See also§ 3.440.) In the case of airplanes equipped withmore than one fuel tank, any tank which is notrequired to feed the engine in all of theconditions specified in this section need beinvestigated only for those flight conditions inwhich it shall be used and the unusable fuelsupply for the particular tank in question shallthen be based on the most critical of thoseconditions which are found to be applicable. Inall such cases, information regarding theconditions under which the full amount of usablefuel in the tank can safely be used shall be madeavailable to the operating personnel by means ofa suitable placard or instruction in the AirplaneFlight Manual.

(b) Upon presentation of the airplane fortest, the applicant shall stipulate the quantity offuel with which he chooses to demonstratecompliance with this provision and shall alsoindicate which of the following conditions ismost critical from the standpoint of establishingthe unusable fuel supply. He shall also indicatethe order in which the other conditions are criticalfrom this standpoint:

(1) Level flight at maximum continuouspower or the power required for level flight at Vc,whichever is less.

(2) Climb at maximum continuous power at

the calculated best angle of climb at minimumweight.

(3) Rapid application of power andsubsequent transition to best rate of climbfollowing a power-off glide at 1.3 Vso.

(4) Sideslips and skids in level flight, climb,and glide under the conditions specified insubparagraphs (1), (2), and (3) of this paragraph,of the greatest severity likely to be encounteredin normal service or in turbulent air.

(c) In the case of utility category airplanes,there shall be no evidence of malfunctioningduring the execution of all approved maneuversincluded in the Airplane Flight Manual. Duringthis test the quantity of fuel in each tank shall notexceed the quantity established as the unusablefuel supply, in accordance with paragraph (b) ofthis section, plus 0.03 gallon for each maximumcontinuous horsepower for which the airplane iscertificated.

(d) In the case of acrobatic categoryairplanes, there shall be no evidence ofmalfunctioning during the execution of allapproved maneuvers included in the AirplaneFlight Manual. During this test the quantity offuel in each tank shall not exceed that specified inparagraph (c) of this section.

(e) If an engine can be supplied with fuelfrom more than one tank, it shall be possible toregain the full power and fuel pressure of thatengine in not more than 10 seconds (for single-engine airplanes) or 20 seconds (for multiengineairplanes) after switching to any full tank afterengine malfunctioning becomes apparent due tothe depletion of the fuel supply in any tank fromwhich the engine can be fed. Compliance withthis provision shall be demonstrated in levelflight.

(f) There shall be no evidence ofmalfunctioning during take-off and climb for 1minute at the calculated attitude of best angle ofclimb at take-off power and minimum weight. Atthe beginning of this test the quantity of fuel ineach tank shall not exceed that specified inparagraph (c) of this section.

§ 3.438 Fuel system hot weather operation.The fuel system shall be so arranged as tominimize the possibility of the formation of vaporlock in the system under all normal conditions ofoperation.

§ 3.439 Flow between interconnectedtanks. In the case of gravity feed systems withtanks who’s outlets are interconnected, it shallnot be possible for fuel to flow between tanks inquantities sufficient to cause an overflow of fuelfrom the tank vent when the airplane is operatedas specified in § 3.437 (a) and the tanks are full.

FUEL TANKS

§ 3.440 General. Fuel tanks shall becapable of withstanding without failure anyvibration, inertia, and fluid and structural loads towhich they may be subjected in operation.Flexible fuel tank liners shall be of an acceptabletype. Integral type fuel tanks shall be providedwith adequate facilities for the inspection andrepair of the tank interior. The total usablecapacity of the fuel tanks shall not be less than 1gallon for each seven maximum continuous ratedhorsepower for which the airplane is certificated.The unusable capacity shall be considered to bethe minimum quantity of fuel which will permitcompliance with the provisions of § 3.437. Thefuel quantity indicator shall be adjusted toaccount for the unusable fuel supply as specifiedin § 3.672. If the unusable fuel supply in any tankexceeds 5 percent of the tank capacity or 1 gallon,whichever is greater, a placard and a suitablenotation in the Airplane Flight Manual shall beprovided to indicate to the flight personnel thatthe fuel remaining in the tank when the quantityindicator reads zero cannot be used safely inflight. The weight of the unusable fuel supplyshall be included in the empty weight of theairplane.

§ 3.441 Fuel tank tests. (a) Fuel tanks shallbe capable of withstanding the followingpressure tests without failure or leakage. Thesepressures may be applied in a manner simulatingthe actual pressure distribution in service:

(1) Conventional metal tanks andnonmetallic tanks whose walls are not supportedby the airplane structure: A pressure of 3.5 psi orthe pressure developed during the maximumultimate acceleration of the airplane with a fulltank, whichever is greater.

(2) Integral tanks: The pressure developedduring the maximum limit acceleration of theairplane with a full tank, simultaneously with theapplication of the critical limit structural loads.

(3) Nonmetallic tanks the walls of which aresupported by the airplane structure: Tanks

constructed of an acceptable basic tank materialand type of construction and with actual orsimulated support conditions shall be subjectedto a pressure of 2 psi for the first tank of aspecific design. Subsequent tanks shall beproduction tested to at least 0.5 psi. Thesupporting structure shall be designed for thecritical loads occurring in the flight or landingstrength conditions combined with the fuelpressure loads resulting from the correspondingaccelerations.

(b) (1) Tanks with large unsupported orunstiffened flat areas shall be capable ofwithstanding the following tests without leakageor failure. The complete tank assembly, togetherwith its supports, shall be subjected to avibration test when mounted in a mannersimulating the actual installation. The tankassembly shall be vibrated for 25 hours at a totalamplitude of not less than 1/32 of an inch whilefilled 2/3 full of water. The frequency of vibrationshall be 90 percent of the maximum continuousrated speed of the engine unless some otherfrequency within the normal operating range ofspeeds of the engine is more critical, in whichcase the latter speed shall be employed and thetime of test shall be adjusted to accomplish thesame number of vibration cycles.

(2) In conjunction with the vibration test,the tank assembly shall be rocked through ananxle of 15° on either side of the horizontal (30°total) about an axis parallel to the axis of thefuselage. The assembly shall be rocked at therate of 16 to 20 complete cycles per minute.

(c) Integral tanks which incorporatemethods of construction and sealing notpreviously substantiated by satisfactory testdata or service experience shall be capable ofwithstanding the vibration test specified inparagraph (b) of this section.

(d) (1) Tanks with nonmetallic liners shall besubjected to the sloshing portion of the testoutlined under paragraph (b) of this section withfuel at room temperature.

(2) In addition, a specimen liner of the samebasic construction as that to be used in theairplane shall, when installed in a suitable testtank, satisfactorily withstand the slosh test withfuel at a temperature of 110°F.

§ 3.442 Fuel tank installation. (a) Themethod of support for tanks shall not be such as

to concentrate the loads resulting from theweight of the fluid in the tanks. Pads shall beprovided to prevent chafing between the tankand its supports. Materials employed forpadding shall be nonabsorbent or shall be treatedto prevent the absorption of fluids. If flexibletank liners are employed, they shall be sosupported that the liner is not required towithstand fluid loads. Interior surfaces ofcompartments for such liners shall be smooth andfree of projections which are apt to cause wear ofthe liner, unless provisions are made forprotection of the liner at such points or unlessthe construction of the liner itself provides suchprotection.

(b) Tank compartments shall be ventilatedand drained to prevent the accumulation ofinflammable fluids or vapors. Compartmentsadjacent of tanks which are an integral part of theairplane structure shall also be ventilated anddrained.

(c) Fuel tanks shall not be located on theengine side of the fire wall. Not less than one-half inch of clear air space shall be providedbetween the fuel tank and the fire wall. Noportion of engine nacelle skin which liesimmediately behind a major air egress openingfrom the engine compartment shall act as the wallof an integral tank. Fuel tanks shall not belocated in personnel compartments, except in thecase of single-engine airplanes. In such casesfuel tanks the capacity of which does not exceed25 gallons may be located in personnelcompartments, if adequate ventilation anddrainage are provided. In all other cases, fueltanks shall be isolated from personnelcompartments by means of fume and fuel proofenclosures.

§ 3.443 Fuel tank expansion space. Fueltanks shall be provided with an expansion spaceof not less than 2 percent of the tank capacity,unless the tank vent discharges clear of theaircraft in which case no expansion space will berequired. It shall not be possible inadvertently tofill the fuel tank expansion space when theairplane is in the normal ground attitude.

§ 3.444 Fuel tank sump. (a) Each tankshall be provided with a drainable sump having acapacity of not less than 0.25 percent of the tankcapacity or 1/16 gallon, whichever is greater. Thesump may be dispensed with if the fuel system isprovided with a sediment bowl which will permit

visual ground inspection for accumulation ofwater or other foreign material. The sedimentbowl shall also be readily accessible for drainage.The capacity of the sediment chamber shall notbe less than 1 ounce per each 20 gallons of thefuel tank capacity.

(b) If a fuel tank sump is provided, thecapacity specified above shall be effective withthe airplane in the normal ground attitude.

(c) If a separate sediment bowl is provided,the fuel tank outlet shall be so located that waterwill drain from all portions of the tank to theoutlet when the airplane is in the ground attitude.

§ 3.445 Fuel tank filler connection. (a)Fuel tank filler connections shall be marked asspecified in § 3.767.

(b) Provision shall be made to prevent theentrance of spilled fuel into the fuel tankcompartment or any portions of the airplane otherthan the tank itself. The filler cap shall provide afuel-tight seal for the main filler opening.However, small openings in the fuel tank cap forventing purposes or to permit passage of a fuelgauge through the cap shall be permissible.

§ 3.446 Fuel tank vents and carburetorvapor vents. (a) Fuel tanks shall be vented fromthe top portion of the expansion space. Ventoutlets shall be so located and constructed as tominimize the possibility of their being obstructedby ice or other foreign matter. The vent shall beso constructed as to preclude the possibility ofsiphoning fuel during normal operation. Thevent shall be of sufficient size to permit the rapidrelief of excessive differences of pressurebetween the interior and exterior of the tank. Airspaces of tanks the outlets of which areinterconnected shall also be interconnected.There shall be no undrainable points in the ventline where moisture is apt to accumulate with theairplane in either the ground or level flightattitude. Vents shall not terminate at pointswhere the discharge of fuel from the vent outletwill constitute a fire hazard or from which fumesmay enter personnel compartments.

(b) Carburetors which are provided withvapor elimination connections shall be providedwith a vent line which will lead vapors back toone of the airplane fuel tanks. If more than onefuel tank is provided and it is necessary to usethese tanks in a definite sequence for any reason,the vapor vent return line shall lead back to the

fuel tank which must be used first unless therelative capacities of the tanks are such thatreturn to another tank is preferable.

§ 3.447-A Fuel tank vents. Provision shall bemade to prevent excessive loss of fuel duringacrobatic maneuvers including short periods ofinverted flight. It shall not be possible for fuel tosiphon from the vent when normal flight has beenresumed after having executed any acrobaticmaneuver for which the airplane is intended.

§ 3.448 Fuel tank outlet. The fuel tankoutlet shall be provided with a screen of from 8 to16 meshes per inch. If a finger strainer is used,the length of the strainer shall not be less than 4times the outlet diameter. The diameter of thestrainer shall not be less than the diameter of thefuel tank outlet. Finger strainers shall beaccessible for inspection and cleaning.

FUEL PUMPS

§ 3.449 Fuel pump and pump installation.(a) If fuel pumps are provided to maintain asupply of fuel to the engine, at least one pumpfor each engine shall be directly driven by theengine. Fuel pumps shall be adequate to meetthe flow requirements of the applicable portionsof §§ 3.433-3.436.

(b) Emergency fuel pumps shall be providedto permit supplying all engines with fuel in caseof the failure of any one engine-driven pump,unless the engine-driven pumps have beenapproved with the engines, in which caseemergency pumps need not be provided.Similarly, if an engine fuel injection pump whichhas been certificated as an integral part of theengine is used, an emergency pump will not berequired. Emergency pumps shall be capable ofcomplying with the same flow requirements as areprescribed for the main pumps. Hand emergencypumps shall not require excessive effort for theircontinued operation at the rate of 60 completecycles (120 single strokes) per minute.Emergency pumps shall be available forimmediate use in case of the failure of any otherpump.

LINES, FITTINGS, AND ACCESSORIES

§ 3.550 Fuel system lines, fittings, andaccessories. Fuel lines shall be installed andsupported in a manner which will preventexcessive vibration and will be adequate towithstand loads due to fuel pressure andaccelerated flight conditions. Lines which are

connected to components of the airplanebetween which relative motion might exist shallincorporate provisions for flexibility. Flexiblehose shall be of an acceptable type.

§ 3.551 Fuel valves. (a) Means shall beprovided to permit the flight personnel to shut offrapidly the flow of fuel to any engine individuallyin flight. Valves provided for this purpose shallbe located on the side of the fire wall most remotefrom the engine.

(b) Shut-off valves shall be so constructedas to make it possible for the flight personnel toreopen the valves rapidly after they have oncebeen closed.

(c) Valves shall be provided with eitherpositive stops or "feel" in the on and offpositions and shall be supported in such amanner that loads resulting from their operationor from accelerated flight conditions are nottransmitted to the lines connected to the valve.Valves shall be so installed that the effect ofgravity and vibration will tend to turn theirhandles to the open rather than the closedposition.

§ 3.552 Fuel strainer. A fuel strainer shallbe provided between the fuel tank outlet and thecarburetor inlet. If an engine-driven fuel pump isprovided, the strainer shall be located betweenthe tank outlet and the engine-driven pump inlet.The strainer shall be accessible for drainage andcleaning, and the strainer screen shall beremovable.

DRAINS AND INSTRUMENTS

§ 3.553 Fuel system drains. Drains shall beprovided to permit safe drainage of the entire fuelsystem and shall incorporate means for locking inthe closed position.

§ 3.554 Fuel system instruments. (See §3.655 and §§ 3.670 through 3.673.)

OIL S YSTEM

§ 3.561 Oil system. Each engine shall beprovided with an independent oil system capableof supplying the engine with an ample quantityof oil at a temperature not exceeding the maximumwhich has been established as safe forcontinuous operation. The oil capacity of thesystem shall not be less than 1 gallon for every25 gallons of fuel capacity. However, in no caseshall the oil capacity be less than 1 gallon foreach 75 maximum continuous horsepower of the

engine(s) involved unless lower quantities can besubstantiated.

§ 3.562 Oil cooling. (See § 3.581 andpertinent sections.)

OIL TANKS

§ 3.563 Oil tanks. Oil tanks shall becapable of withstanding without failure allvibration, inertia, and fluid loads to which theymight be subjected in operation. Flexible oil tankliners shall be of an acceptable type.

§ 3.564 Oil tank tests. Oil tank tests shallbe the same as fuel tank tests (see § 3.441),except as follows:

(a) The 3.5 psi pressure specified in § 3.441(a) shall be 5 pound psi.

(b) In the case of tanks with nonmetabllicliners, the test fluid shall be oil rather than fuel asspecified in § 3.441 (d) and the slosh test on aspecimen liner shall be conducted with oil at atemperature of 250° F.

§ 3.565 Oil tank installation. Oil tankinstallations shall comply with the requirementsof § 3.442 (a) and (b).

§ 3.566 Oil tank expansion space. Oiltanks shall be provided with an expansion spaceof not less than 10 percent of the tank capacity or½ gallon, whichever is greater. it shall not bepossible inadvertently to fill the oil tankexpansion space when the airplane is in thenormal ground attitude.

§ 3.567 Oil tank filler connection. Oil tankfiller connections shall be marked as specified in§ 3.767.

§ 3.568 Oil tank vent. (a) Oil tanks shall bevented to the engine crankcase from the top ofthe expansion space in such a manner that thevent connection is not covered by oil under annormal flight conditions. Oil tank vents shall beso arranged that condensed water vapor whichmight freeze and obstruct the line cannotaccumulate at any point.

(b) Category A. Provision shall be made toprevent hazardous loss of oil during acrobaticmaneuvers including short periods of invertedflight.

§ 3.569 Oil tank outlet. The oil tank outletshall not be enclosed or covered by any screenor other guard which might impede the flow of oil.

The diameter of the oil tank outlet shall not beless than the diameter of the engine oil pumpinlet. (See also § 3.577.)

LINES, FITTINGS, AND ACCESSORIES

§ 3.570 Oil system lines, fittings, andaccessories. Oil lines shall comply with theprovisions of § 3.550, except that the insidediameter of the engine oil inlet and outlet linesshall not be less than the diameter of thecorresponding engine oil pump inlet and outlet.

§ 3.571 Oil valves. (See § 3.637.)

§ 3.572 Oil radiators. Oil radiators andtheir support shall be capable of withstandingwithout failure any vibration, inertia, and oilpressure loads to which they might normally besubjected.

§ 3.573 Oil filters. If the engine is equippedwith an oil filter, the filter shall be constructedand installed in such a manner that completeblocking of the flow through the filter element willnot jeopardize the continued operation of theengine oil supply system.

§ 3.574 Oil system drains. Drains shall beprovided to permit safe drainage of the entire oilsystem and shall incorporate means for positivelocking in the closed position.

§ 3.575 Engine breather lines. (a) Enginebreather lines shall be so arranged thatcondensed water vapor which might freeze andobstruct the line cannot accumulate at any point.Breathers shall discharge in a location which willnot constitute a fire hazard in case foamingoccurs and so that oil emitted from the line willnot impinge upon the pilot’s windshield. Thebreather shall not discharge into the engine airinduction system.

(b) Category A. In the case of acrobatictype airplanes, provision shall be made toprevent excessive loss of oil from the breatherduring acrobatic maneuvers including shortperiods of inverted flight.

§ 3.576 Oil system instruments. See §§3.655, 3.670, 3.671, and 3.674.

§ 3.577 Propeller feathering system. If thepropeller feathering system is dependent uponthe use of the engine oil supply, provision shallbe made to trap a quantity of oil in the tank incase the supply becomes depleted due to failureof any portion of the lubricating system other

than the tank itself. The quantity of oil sotrapped shall be sufficient to accomplish thefeathering operation and shall be available onlyto the feathering pump. The ability of the systemto accomplish feathering when the supply of oilhas fallen to the above level shall bedemonstrated.

COOLING

§ 3.581 General. The power-plant coolingprovisions shall be capable of maintaining thetemperatures of all power-plant components,engine parts, and engine fluids (oil and coolant),at or below the maximum established safe valuesunder critical conditions of ground and flightoperation.

TESTS

§ 3.582 Cooling tests. Compliance with theprovisions of § 3.581 shall be demonstratedunder critical ground, water, and flight operatingconditions. If the tests are conducted underconditions which deviate from the highestanticipated summer air temperature (see § 3.583),the recorded power-plant temperatures shall becorrected in accordance with the provisions of§§ 3.584 and 3.585. The corrected temperaturesdetermined in this manner shall not exceed themaximum established safe values. The fuel usedduring the cooling tests shall be of the minimumoctane number approved for the enginesinvolved, and the mixture setting shall be thoseappropriate to the operating conditions. The testprocedures shall be as outlined in §§ 3.586 and3.587.

§ 3.583 Maximum anticipated summer airtemperatures. The maximum anticipated summerair temperature shall be considered to be 100° F,at sea level and to decrease from this value at therate of 3.6° F, per thousand feet of altitude abovesea level.

§ 3.584 Correction factor for cylinderhead, oil inlet, carburetor air, and enginecoolant inlet temperatures. These temperaturesshall be corrected by adding the differencebetween the maximum anticipated summer airtemperature and the temperature of the ambientair at the time of the first occurrence of maximumhead, air, oil, or coolant temperature recordedduring the cooling test.

§ 3.585 Correction factor for cylinderbarrel temperatures. Cylinder barreltemperatures shall be corrected by adding 0.7 of

the difference between the maximum anticipatedsummer air temperature and the temperature ofthe ambient air at the first occurrence of themaximum cylinder barrel temperature recordedduring the cooling test.

§ 3.586 Cooling test procedure for single-engine airplanes. This test shall be conductedby stabilizing engine temperatures in flight andthen starting at the lowest practicable altitudeand climbing for 1 minute at take-off power. Atthe end of 1 minute, the climb shall be continuedat maximum continuous power until at least 5minutes after the occurrence of the highesttemperature recorded. The climb shall not beconducted at a speed greater than the best rate-of-climb speed with maximum continuous powerunless:

(a) The slope of the flight path at the speedchosen for the cooling test is equal to or greaterthan the minimum required angle of climb (see §3.85 (a)), and

(b) A cylinder head temperature indicator isprovided as specified in § 3.675.

§ 3.587 Cooling test procedure formultiengine airplanes—(a) Airplanes whichmeet the minimum one-engine-inoperative climbperformance specified in § 3.85 (b). The enginecooling test for these airplanes shall beconducted with the airplane in the configurationspecified in § 3.85 (b), except that the operatingengine(s) shall be operated at maximumcontinuous power or at full throttle when abovethe critical altitude. After stabilizing temperaturesin flight, the climb shall be started at the lower ofthe two following altitudes and shall becontinued until at least 5 minutes after thehighest temperature has been recorded:

(1) 1,000 feet below the engine criticalaltitude or at the lowest practicable altitude(when applicable).

(2) 1,000 feet below the altitude at which thesingle-engine-inoperative rate of climb is 0.02Vso

2.

The climb shall be conducted at a speed notin excess of the highest speed at whichcompliance with the climb requirement of § 3.85(b) can be shown. However, if the speed usedexceeds the speed for best rate of climb with oneengine inoperative, a cylinder head temperatureindicator shall be provided as specified in § 3.675.

(b) Airplanes which cannot meet theminimum one-engine-inoperative climbperformance specified in § 3.85 (b). The enginecooling test for these airplanes shall be the sameas in paragraph (a) of this section, except thatafter stabilizing temperatures in flight, the climb(or descent, in the case of airplanes with zero ornegative one-engine-inoperative rate of climb)shall be commenced at as near sea level aspracticable and shall be conducted at the bestrate-of-climb speed (or the speed of minimum rateof descent, in the case of airplanes with zero ornegative one-engine-inoperative rate of climb).

LIQUID COOLING S YSTEMS

§ 3.588 Independent systems. Each liquidcooled engine shall be provided with anindependent cooling system. The coolingsystem shall be so arranged that no air or vaporcan be trapped in any portion of the system,except the expansion tank, either during filling orduring operation.

§ 3.589 Coolant tank. A coolant tank shallbe provided. The tank capacity shall not be lessthan 1 gallon plus 10 percent of the coolingsystem capacity. Coolant tanks shall be capableof withstanding without failure all vibration,inertia, and fluid loads to which they may besubjected in operation. Coolant tanks shall beprovided with an expansion space of not lessthan 10 percent of the total cooling systemcapacity. It shall not be possible inadvertently tofill the expansion space with the airplane in thenormal ground attitude.

§ 3.590 Coolant tank tests. Coolant tanktests shall be the same as fuel tank tests (see §3.441), except as follows:

(a) The 3.5 pounds per square inchpressure test of § 3.441 (a) shall be replaced bythe sum of the pressure developed during themaximum ultimate acceleration with a full tank or apressure of 3.5 pounds per square inch,whichever is greater, plus the maximum workingpressure of the system.

(b) In the case of tanks with nonmetallicliners, the test fluid shall be coolant rather thanfuel as specified in § 3.441 (d), and the slosh teston a specimen liner shall be conducted withcoolant at operating temperature.

§ 3.591 Coolant tank installation. Coolanttanks shall be supported in a manner so as todistribute the tank loads over a large portion of

the tank surface. Pads shall be provided toprevent chafing between the tank and thesupport. Material used for padding shall benonabsorbent or shall be treated to prevent theabsorption of inflammable fluids.

§ 3.592 Coolant tank filler connection.Coolant tank filler connections shall be marked asspecified in § 3.767. Provisions shall be made toprevent the entrance of spilled coolant into thecoolant tank compartment or any portions of theairplane other than the tank itself. Recessedcoolant filler connections shall be drained andthe drain shall discharge clear of all portions ofthe airplane.

§ 3.593 Coolant lines, fittings, andaccessories. Coolant lines shall comply with theprovisions of § 3.550, except that the insidediameter of the engine coolant inlet and outletlines shall not be less than the diameter of thecorresponding engine inlet and outletconnections.

§ 3.594 Coolant radiators. Coolantradiators shall be capable of withstandingwithout failure any vibration, inertia, and coolantpressure loads to which they may normally besubjected. Radiators shall be supported in amanner which will permit expansion due tooperating temperatures and prevent thetransmittal of harmful vibration to the radiator. Ifthe coolant employed is inflammable, the airintake duct to the coolant radiator shall be solocated that flames issuing from the nacelle incase of fire cannot impinge upon the radiator.

§ 3.595 Cooling system drains. One ormore drains shall be provided to permit drainageof the entire cooling system, including thecoolant tank, radiator, and the engine, when theairplane is in the normal ground attitude. Drainsshall discharge clear of all portions of the airplaneand shall be provided with means for positivelylocking the drain in the closed position. Coolingsystem drains shall be accessible.

§ 3.596 Cooling system instruments. See§§ 3.655, 3.670, and 3.671.

INDUCTION S YSTEM

§ 3.605 General. (a) The engine airinduction system shall permit supplying anadequate quantity of air to the engine under allconditions of operation.

(b) Each engine shall be provided with at

least two separate air intake sources, except thatin the case of an engine equipped with a fuelinjector only one air intake source need beprovided, if the air intake, opening, or passage isunobstructed by a screen, filter, or other part onwhich ice might form and so restrict the air flowas to affect adversely engine operation. Primaryand alternate air intakes may open within thecowling only if that portion of the cowling isisolated from the engine accessory section bymeans of a fireproof diaphragm. Alternate airintakes shall be located in a sheltered position.Supplying air to the engine through the alternateair intake system or the carburetor air preheatershall not result in the loss of excessive power inaddition to the power lost due to the rise in thetemperature of the air.

§ 3.606 Induction system de-icing and anti-icing provisions. The engine air inductionsystem shall incorporate means for theprevention and elimination of ice accumulationsin accordance with the provisions in this section.It shall be demonstrated that compliance with theprovisions outlined in the following paragraphscan be accomplished when the airplane isoperating in air at a temperature of 30° F, whenthe air is free of visible moisture.

(a) Airplanes equipped with sea levelengines employing conventional venturicarburetors shall be provided with a preheatercapable of providing a heat rise of 90° F. whenthe engine is operating at 75 percent of itsmaximum continuous power.

(b) Airplanes equipped with altitudeengines employing conventional venturicarburetors shall be provided with a preheatercapable of providing a heat rise of 120° F. whenthe engine is operating at 75 percent of itsmaximum continuous power.

(c) Airplanes equipped with altitudeengines employing carburetors which embodyfeatures tending to reduce the possibility of iceformation shall be provided with a preheatercapable of providing a heat rise of 100° F. whenthe engine is operating at 60 percent of itsmaximum continuous power. However, thepreheater need not provide a heat rise in excessof 40°F. if a fluid de-icing system complying withthe provisions of §§ 3.607-3.609 is also installed.

§ 3.607 Carburetor de-icing fluid flow rate.The system shall be capable of providing each

engine with a rate of fluid flow, expressed inpounds per hour, of not less than 2.5 multipliedby the square root of the maximum continuouspower of the engine. This flow shall be availableto all engines simultaneously. The fluid shall beintroduced into the air induction system at apoint close to, and upstream from, the carburetor.The fluid shall be introduced in a manner toassure its equal distribution over the entire crosssection of the induction system air passages.

§ 3.608 Carburetor fluid de-icing systemcapacity. The fluid de-icing system capacityshall not be less than that required to providefluid at the rate specified in § 3.607 for a timeequal to 3 percent of the maximum endurance ofthe airplane. However, the capacity need not inany case exceed that required for 2 hours ofoperation nor shall it be less than that requiredfor 20 minutes of operation at the above flow rate.If the available preheat exceeds 50° F. but is lessthan 100° F., it shall be permissible to decreasethe capacity of the system in proportion to theheat rise available in excess of 50° F.

§ 3.609 Carburetor fluid de-icing systemdetail design. Carburetor fluid de-icing systemsshall comply with provisions for the design offuel systems, except as specified in §§ 3.607 and3.608, unless such provisions are manifestly inapplicable.

§ 3.610 Carburetor air preheater design.Means shall be provided to assure adequateventilation of the carburetor air preheater whenthe engine is being operated in cold air. Thepreheater shall be constructed in such a manneras to permit inspection of exhaust manifold partswhich it surrounds and also to permit inspectionof critical portions of the preheater itself.

§ 3.611 Induction system ducts. Inductionsystem ducts shall be provided with drains whichwill prevent the accumulation of fuel or moisturein all normal ground and flight attitudes. Noopen drains shall be located on the pressure sideof turbo-supercharger installations. Drains shallnot discharge in a location which will constitute afire hazard. Ducts which are connected tocomponents of the airplane between whichrelative motion may exist shall incorporateprovisions for flexibility.

§ 3.612 Induction system screens. Ifinduction system screens are employed, theyshall be located upstream from the carburetor. It

shall not be possible for fuel to impinge upon thescreen. Screens shall not be located in portionsof the induction system which constitute theonly passage through which air can reach theengine, unless the available preheat is 100° F, orover and the screen is so located that it can bede-iced by the application of heated air. De-icingof screens by means of alcohol in lieu of heatedair shall not be acceptable.

EXHAUST S YSTEM

§ 3.615 General. (a) The exhaust systemshall be constructed and arranged in such amanner as to assure the safe disposal of exhaustgases without the existence of a hazard of fire orcarbon monoxide contamination of air inpersonnel compartments.

(b) Unless suitable precautions are taken,exhaust system parts shall not be located in closeproximity to portions of any systems carryinginflammable fluids or vapors nor shall they belocated under portions of such systems whichmay be subject to leakage. All exhaust systemcomponents shall be separated from adjacentinflammable portions of the airplane which areoutside the engine compartment by means offireproof shields. Exhaust gases shall not bedischarged at a location which will cause a glareseriously affecting pilot visibility at night, norshall they discharge within dangerous proximityof any fuel or oil system drains. All exhaustsystem components shall be ventilated toprevent the existence of points of excessivelyhigh temperature.

§ 3.616 Exhaust manifold. Exhaustmanifolds shall be made of fireproof, corrosion-resistant materials, and shall incorporateprovisions to prevent failure due to theirexpansion when heated to operatingtemperatures. Exhaust manifolds shall besupported in a manner adequate to withstand allvibration and inertia loads to which they might besubjected in operation. Portions of the manifoldwhich are connected to components betweenwhich relative motion might exist shallincorporate provisions for flexibility.

§ 3.617 Exhaust heat exchangers. (a)Exhaust heat exchanges shall be constructed andinstalled in such a manner as to assure theirability to withstand without failure all vibration,inertia, and other loads to which they mightnormally be subjected. Heat exchangers shall beconstructed of materials which are suitable for

continued operation at high temperatures andwhich are adequately resistant to corrosion dueto products contained in exhaust gases.

(b) Provisions shall be made for theinspection of all critical portions of exhaust heatexchangers, particularly if a welded constructionis employed. Heat exchangers shall be ventilatedunder all conditions in which they are subject tocontact with exhaust gases.

§ 3.618 Exhaust heat exchangers used inventilating air heating systems. Heat exchangersof this type shall be so constructed as topreclude the possibility of exhaust gases enteringthe ventilating air.

FIRE WALL AND COWLING

§ 3.623 Fire walls. All engines, auxiliarypower units, fuel burning heaters, and othercombustion equipment which are intended foroperation in flight shall be isolated from theremainder of the airplane by means of fire walls,or shrouds, or other equivalent means.

§ 3.624 Fire wall construction. (a) Firewalls and shrouds shall be constructed in such amanner that no hazardous quantity of air, fluids,or flame can pass from the engine compartment toother portions of the airplane. All openings inthe fire wall or shroud shall be sealed with close-fitting fireproof grommets, bushings, or fire-wallfittings.

(b) Fire walls and shrouds shall beconstructed of fireproof material and shall beprotected against corrosion. The followingmaterials have been found to comply with thisrequirement:

(1) Heat- and corrosion-resistant steel 0.015inch thick,

(2) Low carbon steel, suitably protectedagainst corrosion, 0.018 inch thick.

§ 3.625 Cowling. (a) Cowling shall beconstructed and supported in such a manner asto be capable of resisting all vibration, inertia,and air loads to which it may normally besubjected. Provision shall be made to permitrapid and complete drainage of all portions of thecowling in all normal ground and flight attitudes.Drains shall not discharge in locationsconstituting a fire hazard.

(b) Cowling shall be constructed of fire-

resistant material. All portions of the airplanelying behind openings in the engine compartmentcowling shall also be constructed of fire-resistantmaterials for a distance of at least 24 inches aft ofsuch openings. Portions of cowling which aresubjected to high temperatures due to proximityto exhaust system ports or exhaust gasimpingement shall be constructed of fireproofmaterial.

POWER-PLANT CONTROLS AND ACCESSORIES

CONTROLS

§ 3.627 Power-plant controls. Power-plantcontrols shall comply with the provisions of §§3.384 and 3.759. Controls shall maintain anynecessary position without constant attention bythe flight personnel and shall not tend to creepdue to control loads or vibration. Flexiblecontrols shall be of an acceptable type. Controlsshall have adequate strength and rigidity towithstand operating loads without failure orexcessive deflection.

§ 3.628 Throttle controls. A throttlecontrol shall be provided to give independentcontrol for each engine. Throttle controls shallafford a positive and immediately responsivemeans of controlling the engine(s). Throttlecontrols shall be grouped and arranged in such amanner as to permit separate control of eachengine and also simultaneous control of allengines.

§ 3.629 Ignition switches. Ignitionswitches shall provide control for each ignitioncircuit on each engine. It shall be possible toshut off quickly all ignition on multiengineairplanes, either by grouping of the individualswitches or by providing a master ignitioncontrol.

§ 3.630 Mixture controls. If mixturecontrols are provided, a separate control shall beprovided for each engine. The controls shall begrouped and arranged in such a manner as topermit both separate and simultaneous control ofall engines.

§ 3.631 Propeller speed and pitch controls.(See also § 3.421 (a).) If propeller speed or pitchcontrols are provided, the controls shall begrouped and arranged in such a manner as topermit control of all propellers, both separatelyand together. The controls shall permit readysynchronization of all propellers on multiengine

airplanes.

§ 3.632 Propeller feathering controls. Ifpropeller feathering controls are provided, aseparate control shall be provided for eachpropeller. Propeller feathering controls shall beprovided with means to prevent inadvertentoperation.

§ 3.633 Fuel system controls. Fuel systemcontrols shall comply with requirements of §3.551 (c).

§ 3.634 Carburetor air preheat controls.Separate control shall be provided to regulate thetemperature of the carburetor air for each engine.

ACCESSORIES

§ 3.635 Power-plant accessories. Engine-driven accessories shall be of a type satisfactoryfor installation on the engine involved and shallutilize the provisions made on the engine for themounting of such units. Items of electricalequipment subject to arcing or sparking shall beinstalled so as to minimize the possibility of theircontact with any inflammable fluids or vaporswhich might be present in a free state.

§ 3.636 Engine battery ignition systems.(a) Battery ignition systems shall besupplemented with a generator which isautomatically made available as an alternatesource of electrical energy to permit continuedengine operation in the event of the depletion ofany battery.

(b) The capacity of batteries and generatorsshall be sufficient to meet the simultaneousdemands of the engine ignition system and thegreatest demands of any of the airplane’selectrical system components which may drawelectrical energy from the same source.Consideration shall be given to the condition ofan inoperative generator, and to the condition ofa completely depleted battery when the generatoris running at its normal operating speed. If onlyone battery is provided, consideration shall alsobe given to the condition in which the battery iscompletely depleted and the generator isoperating at idling speed.

(c) Means shall be provided to warn theappropriate flight personnel if malfunctioning ofany part of the electrical system is causing thecontinuous discharging of a battery used forengine ignition. (See § 3.629 for ignitionswitches.)

POWER-PLANT FIRE PROTECTION

§ 3.637 Power-plant fire protection.Suitable means shall be provided to shut off theflow in all lines carrying inflammable fluids intothe engine compartment.

S UBPART F - EQUIPMENT

§ 3.651 General. The equipment specifiedin § 3.655 shall be the minimum installed when theairplane is submitted to determine its compliancewith the airworthiness requirements. Suchadditional equipment as is necessary for aspecific type of operation is specified in otherpertinent parts of the Civil Air Regulations, but,where necessary, its installation and that of theitems mentioned in § 3.655 is covered herein.

§ 3.652 Functional and installationalrequirements. Each item of equipment which isessential to the safe operation of the airplaneshall be found by the Administrator to performadequately the functions for which it is to beused, shall function properly when installed, andshall be adequately labeled as to itsidentification, function, operational limitations, orany combination of these, whichever isapplicable. Items of equipment for which typecertification is required shall have beencertificated in accordance with the provisions ofPart 15 of this chapter (or previous regulations)and such other parts as may be applicable.

BASIC EQUIPMENT

§ 3.655 Required basic equipment. Thefollowing table shows the basic equipment itemsrequired for type and airworthiness certificationof an airplane:

(a) Flight and navigational instruments.(1) Air-speed indicator (see § 3.663).

(2) Altimeter.

(3) Magnetic direction indicator (see §3.666)

(b) Power-plant instruments—(1) For eachengine or tank. (i) Fuel quantity indicator (see §3.672).

(ii) Oil pressure indicator.

(iii) Oil temperature indicator.

(iv) Tachometer.

(2) For each engine or tank (if required inreference section). (i) Carburetor air temperatureindicator (see § 3.676).

(ii) Coolant temperature indicator (if liquid-cooled engines used).

(iii) Cylinder head temperature indicator (see§ 3.675).

(iv) Fuel pressure indicator (if pump-fedengines used).

(v) Manifold pressure indicator (if altitudeengines used).

(vi) Oil quantity indicator (see § 3.674).

(c) Electrical equipment (if required byreference section). (1) Master switcharrangement (see § 3.688).

(2) Adequate source(s) of electrical energy(see §§ 3.682 and 3.685).

(3) Electrical protective devices (see §3.690).

(d) Miscellaneous equipment. (1)Certificated safety belts for all occupants (seePart 15 of this chapter).

(2) Airplane Flight Manual (see § 3.777).

INSTRUMENTS; INSTALLATION

GENERAL

§ 3.661 Arrangement and visibility ofinstrument installations. (a) Flight, navigation,and power-plant instruments for use by eachpilot shall be easily visible to him.

(b) On multiengine airplanes, identicalpower-plant instruments for the several enginesshall be so located as to, prevent any confusionas to the engines to which they relate.

§ 3.662 Instrument panel vibrationcharacteristics. Vibration characteristics of theinstrument panel shall not be such as to impairthe accuracy of the instruments or to causedamage to them.

FLIGHT AND NAVIGATIONAL INSTRUMENTS

§ 3.663 Air-speed indicating system. Thissystem shall be so installed that the air-speedindicator shall indicate true air speed at sea levelunder standard conditions to within an allowableinstallational error of not more than plus or minus3 percent of the calibrated air speed or 5 miles perhour, whichever is greater, throughout theoperating range of the airplane with flaps up fromVc to 1.3 Vs1 and with flaps at 1.3 Vs1. Thecalibration shall be made in flight.

§ 3.664 Air-speed indicator-marking. The

air-speed indicator shall be marked as specified in§ 3.757.

§ 3.665 Static air vent system. Allinstruments provided with static air caseconnections shall be so vented that the influenceof airplane speed, the opening and closing ofwindows, air-flow variation, moisture, or otherforeign matter will not seriously affect theiraccuracy.

§ 3.666 Magnetic direction indicator. Themagnetic direction indicator shall be so installedthat its accuracy shall not be excessively affectedby the airplane’s vibration or magnetic fields.After the direction indicator has beencompensated, the installation shall be such thatthe deviation in level flight does not exceed 10degrees on any heading. A suitable calibrationplacard shall be provided as specified in § 3.758.

§ 3.667 Automatic pilot system. If anautomatic pilot system is installed:

(a) The actuating (servo) devices shall beof such design that they can, when necessary, bepositively disengaged from operating the controlsystem or be overpowered by the human pilot toenable him to maintain satisfactory control of theairplane.

(b) A satisfactory means shall be providedto indicate readily to the pilot the alignment ofthe actuating device in relation to the controlsystem which it operates, except when automaticsynchronization is provided.

(c) The manually operated control(s) for thesystem’s operation shall be readily accessible tothe pilot.

(d) The automatic pilot system shall be ofsuch design and so adjusted that it cannotproduce loads in the control system and surfacesgreater than those for which they were designed.

§ 3.668 Gyroscopic indicators (air-driventype). All air-driven gyroscopic instrumentsinstalled in airplanes which are certificated forinstrument flight operations shall derive theirenergy from a reliable suction source of sufficientcapacity to maintain their required accuracy at allspeeds above the best rate-of-climb speed. Inaddition the system shall be so installed as topreclude malfunctioning due to rain, oil, or otherdetrimental elements. On multiengine airplanes,the following detail requirements shall beapplicable:

(a) Two sources actuated by separatemeans shall be provided, either one of whichshall be of sufficient capacity to operate all of theair-driven gyroscopic instruments with which theairplane is equipped, with the airplane in normalcruising attitude at 65 percent maximumcontinuous power.

(b) A suitable means shall be provided inthe attendant installation where the source linesconnect into a common line to select eithersuction air source for the proper functioning ofthe instruments should failure of one source or abreakage of one source line occur. When anautomatic means to permit simultaneous air flowis provided in the system, a suitable method formaintaining suction shall be provided. In orderto indicate which source of energy has failed, avisual means shall be provided to indicate thiscondition to the flight crew.

§ 3.669 Suction gauge. A suction gaugeshall be provided and so installed as to indicatereadily to the flight crew while in flight thesuction in inches of mercury which is beingapplied to the air-driven types of gyroscopicinstruments. This gauge shall be connected tothe instruments by a suitable system.

POWER-PLANT INSTRUMENTS

§ 3.670 Operational markings.Instruments shall be marked as specified in §3.759.

§ 3.671 Instrument lines. Power-plantinstrument lines shall comply with the provisionsof § 3.550. In addition, instrument lines carryinginflammable fluids or gases under pressure shallbe provided with restricted orifices or othersafety devices at the source of the pressure toprevent escape of excessive fluid or gas in caseof line failure.

§ 3.672 Fuel quantity indicator. Meansshall be provided to indicate to the flightpersonnel the quantity of fuel in each tank duringflight. Tanks, the outlets and air spaces of whichare interconnected, may be considered as onetank and need not be provided with separateindicators. Exposed sight gauges shall be soinstalled and guarded as to preclude thepossibility of breakage or damage. Fuel quantityindicators shall be calibrated to read zero duringlevel flight when the quantity of fuel remaining inthe tank is equal to the unusable fuel supply asdefined by § 3.437.

[12 F.R. 3438]

§ 3.673 Fuel flowmeter system. When afuel flowmeter system is installed in the fuelline(s), the metering component shall be of suchdesign as to include a suitable means forbypassing the fuel supply in the event thatmalfunctioning of the metering component offersa severe restriction to fuel flow.

§ 3.674 Oil quantity indicator. Groundmeans, such as a slick gauge, shall be providedto indicate the quantity of oil in each tank. If anoil transfer system or a reserve oil supply systemis installed, means shall be provided to indicateto the flight personnel during flight the quantityof oil in each tank.

§ 3.675 Cylinder head temperatureindicating system for air-cooled engines. Acylinder head temperature indicator shall beprovided for each engine on airplanes equippedwith cowl flaps. In the case of airplanes which donot have cowl flaps, an indicator shall beprovided if compliance with the provisions of §3.581 is demonstrated at a speed in excess of thespeed of best rate of climb.

§ 3.676 Carburetor air temperatureindicating system. A carburetor air temperatureindicating system shall be provided for eachaltitude engine equipped with a preheater whichis capable of providing a heat rise in excess of60°F.

ELECTRICAL S YSTEMS AND EQUIPMENT

§ 3.681 Installation. (a) Electrical systemsin airplanes shall be free from hazards inthemselves, in their method of operation, and intheir effects on other parts of the airplane.Electrical equipment shall be of a type and designadequate for the use intended. Electrical systemsshall be installed in such a manner that they aresuitably protected from fuel, oil, water, otherdetrimental substances, and mechanical damage.

(b) Items of electrical equipment requiredfor a specific type of operation are listed in otherpertinent parts of the Civil Air Regulations.

BATTERIES

§ 3.682 Batteries. When an item ofelectrical equipment which is essential to the safeoperation of the airplane is installed, the battery

required shall have sufficient capacity to supplythe electrical power necessary for dependableoperation of the connected electrical equipment.

§ 3.683 Protection against acid. Ifbatteries are of such a type that corrosivesubstance may escape during servicing or flight,means such as a completely enclosedcompartment shall be provided to prevent suchsubstances from coming in contact with otherparts of the airplane which are essential to safeoperation. Batteries shall be accessible forservicing and inspection on the ground.

§ 3.684 Battery vents. The batterycontainer or compartment shall be vented in suchmanner that gases released by the battery arecarried outside the airplane.

GENERATORS

§ 3.685 Generator. Generators shall becapable of delivering their continuous ratedpower.

§ 3.686 Generator controls. Generatorvoltage control equipment shall be capable ofdependably regulating the generator outputwithin rated limits.

§ 3.687 Reverse current cut-out. Agenerator reverse current cut-out shalldisconnect the generator from the battery andother generators when the generator isdeveloping a voltage of such value that currentsufficient to cause malfunctioning can flow intothe generator.

MASTER S WITCH

§ 3.688 Arrangement. If electricalequipment is installed, a master switcharrangement shall be provided which willdisconnect all sources of electrical power fromthe main distribution system at a point adjacentto the power sources.

§ 3.689 Master switch installation. Themaster switch or its controls shall be so installedthat it is easily discernible and accessible to amember of the crew in flight.

PROTECTIVE DEVICES

§ 3.690 Fuses or circuit breakers. Ifelectrical equipment is installed, protectivedevices (fuses or circuit breakers) shall beinstalled in the circuits to all electrical equipment,except that such items need not be installed inthe main circuits of starter motors or in other

circuits where no hazard is presented by theiromission.

§ 3.691 Protective devices installation.Protective devices in circuits essential to safetyin flight shall be so located and identified thatfuses may be replaced or circuit breakers resetreadily in flight.

§ 3.692 Square fuses. If fuses are used, onespare of each rating or 50 percent spare fuses ofeach rating, whichever is greater, shall beprovided.

ELECTRIC CABLES

§ 3.693 Electric cables. If electricalequipment is installed, the connecting cablesused shall be in accordance with recognizedstandards for electric cable of a slow burningtype and of suitable capacity.

S WITCHES

§ 3.694 Switches. Switches shall becapable of carrying their rated current and shallbe of such construction that there is sufficientdistance or insulating material between currentcarrying parts and the housing so that vibrationin flight will not cause shorting.

§ 3.695 Switch installation. Switches shallbe so installed as to be readily accessible to theappropriate crew member and shall be suitablylabeled as to operation and the circuit controlled.

INSTRUMENT LIGHTS

§ 3.696 Instrument lights. If instrumentlights are required, they shall be of suchconstruction that there is sufficient distance orinsulating material between current carrying partsand the housing so that vibration in flight will notcause shorting. They shall provide sufficientillumination to make all instruments and controlseasily readable and discernible, respectively.

§ 3.697 Instrument light installation.Instrument lights shall be installed in such amanner that their direct rays are shielded from thepilot’s eyes. Direct rays shall not be reflectedfrom the windshield or other surfaces into thepilot’s eyes.

LANDING LIGHTS

§ 3.698 Landing lights. If landing lights are

installed, they shall be of an acceptable type.

§ 3.699 Landing light installation.Landing lights shall be so installed that there isno dangerous glare visible to the pilot and alsoso that the pilot is not seriously affected byhalation. They shall be installed at such alocation that they provide adequate illuminationfor night landing.

POSITION LIGHTS

§ 3.700 Type. If position lights areinstalled, they shall be of a type certificated inaccordance with Part 15 of this chapter, or shallcomply with the pertinent provisions of that part.

§ 3.701 Forward position lightinstallation. Forward position lights shall be soinstalled that, with the airplane in normal flyingposition, the red light is displayed on the left sideand the green light on the right side, eachshowing unbroken light between two verticalplanes the dihedral angle of which is 100 degreeswhen measured to the left and right, respectively,of the airplane from dead ahead. The lights shallbe spaced laterally as far apart as practicable.

§ 3.702 Rear position light installation.The rear position light shall be mounted as far aftas practicable and so installed that unbrokenlight is directed symmetrically aft in such amanner that the axis of the maximum cone ofillumination is parallel to the flight path. Inaddition, the intersection of the two planesforming dihedral angle A given in Part 15 of thischapter shall be vertical.

§ 3.703 Flashing rear position lights. Ifred and white flashing lights are used, in additionto meeting the installation requirements in §3.702, they shall be located close together.

ANCHOR LIGHTS

§ 3.704 Anchor light. When an anchorlight is required for seaplanes and amphibians, atleast one light shall be provided and it shall becapable of showing a white light for at least 2miles at night under clear atmosphere conditions.

§ 3.705 Anchor light installation. Anchorlights shall be so installed that they will show themaximum unbroken light practicable when theairplane is moored or drifting on the water.Externally hung lights are permitted.

S AFETY EQUIPMENT; INSTALLATION

§ 3.711 Marking. Required safetyequipment which the crew is expected to operateat a time of emergency, such as flares andautomatic life raft releases, shall be readilyaccessible and plainly marked as to its method ofoperation. When such equipment is carried inlockers, compartments, or other storage places,such storage places shall be marked for thebenefit of passengers and crew.

§ 3.712 De-icers. When pneumatic de-icers are installed, the installation shall be inaccordance with approved data. Positive meansshall be provided for the deflation of thepneumatic boots.

§ 3.713 Flare requirements. Whenparachute flares are required, they shall be of atype certificated in accordance with Part 15 ofthis chapter.

§ 3.714 Flare installation. Parachute flaresshall be releasable from the pilot compartmentand so installed that danger of accidentaldischarge is reduced to a minimum. Theinstallation shall be demonstrated in flight toeject flares satisfactorily, except in those caseswhere inspection indicates a ground test will beadequate. If the flares are ejected so that recoilloads are involved , structural provisions forsuch loads shall be made.

§ 3.715 Safety belts. Safety belts shall be ofa type certificated in accordance with Part 15 ofthis chapter. They shall be so attached that nopart of the anchorage will fail at a lower load thanspecified in § 3.386.

EMERGENCY FLOTATION AND S IGNALING EQUIPMENT

§ 3.716 Rafts and life preservers. Anapproved life raft or approved life preserver,when required by other parts of the Civil AirRegulations, is one approved by either theAdministrator, the Bureau of Marine Inspectionand Navigation, the United States Army AirForces, or the Bureau of Aeronautics, NavyDepartment.

§ 3.717 Installation. When suchemergency equipment is required, it shall be soinstalled as to be readily available to the crew andpassengers. Rafts released automatically or bythe pilot shall be attached to the airplane bymeans of a line to keep them adjacent to theairplane. The strength of the line shall be suchthat it will break before submerging the empty

raft.

§ 3.718 Signaling device. Signalingdevices, when required by other parts of the CivilAir Regulations, shall be accessible, functionsatisfactorily, and be free from any hazard in theiroperation.

RADIO EQUIPMENT; INSTALLATION

§ 3.721 General. Radio equipment andinstallations in the airplane shall be free fromhazards in themselves, in their method ofoperation, and in their effects on theircomponents of the airplane.

MISCELLANEOUS EQUIPMENT; INSTALLATION

§ 3.725 Accessories for multiengineairplanes. Engine driven accessories essential tothe safe operation of the airplane shall be sodistributed among two or more engines that thefailure of any one engine will not impair the safeoperation of the airplane by the malfunctioning ofthese accessories.

HYDRAULIC S YSTEMS

§ 3.726 General. Hydraulic systems andelements shall be so designed as to withstand,without exceeding the yield point, any structuralloads which might be imposed in addition to thehydraulic loads.

§ 3.727 Tests. Hydraulic systems shall besubstantiated by proof pressure tests. Whenproof tested, no part of the hydraulic systemshall fail, malfunction, or experience a permanentset. The proof load of any system shall be 15times the maximum operating pressure of thatsystem.

§ 3.728 Accumulators. Hydraulicaccumulators or pressurized reservoirs shall notbe installed on the engine side of the fire wall,except when they form an integral part of theengine or propeller.

S UBPART G——OPERATING LIMITATIONS AND

INFORMATION

§ 3.735 General. Means shall be providedto inform adequately the pilot and otherappropriate crew members of all operatinglimitations upon which the type design is based.Any other information concerning the airplanefound by the Administrator to be necessary forsafety during its operation shall also be madeavailable to the crew. (See §§ 3.755 and 3.777.)

LIMITATIONS

§ 3.737 Limitations. The operatinglimitations specified in §§ 3.738-3.750 and anysimilar limitations shall be established for anyairplane and made available to the operator asfurther described in §§ 3.755-3.780, unless itsdesign is such that they are unnecessary for safeoperation.

AIR S PEED

§ 3.738 Air speed. Air-speed limitationsshall be established as set forth in §§ 3.739-3.743.

§ 3.739 Never-exceed speed (Vne). Thisspeed shall not exceed the lesser of the following:

(a) 0.9 Vd chosen in accordance with §3.184.

(b) 0.9 times the maximum speeddemonstrated in accordance with § 3.159, butshall not be less than 0.9 times the minimum valueof Vd permitted by § 3.184.

§ 3.740 Maximum structural cruising speed(Vno). This operating limitation shall be:

(a) Not greater than Vc chosen inaccordance with § 3.184.

(b) Not greater than 0.89 times Vneestablished under § 3.739.

(c) Not less than the minimum Vc permittedin § 3.184.

§ 3.741 Maneuvering speed (Vp). (See §3.184.)

§ 3.742 Flaps-extended speed (Vfe). (a)This speed shall not exceed the lesser of thefollowing:

(1) The design flap speed, Vf chosen inaccordance with § 3.190.

(2) The design flap speed chosen inaccordance with § 3.223, but shall not be lessthan the minimum value of design flap speedpermitted in §§ 3.190 and 3.223.

(b) Additional combinations of flap setting,air speed, and engine power may be established,provided the structure has been proven for thecorresponding design conditions.

§ 3.743 Minimum control speed (Vmc).(See § 3.111.)

POWER PLANT

§ 3.744 Power plant. The power plantlimitations in §§ 3.745 through 3.747 shall beestablished and shall not exceed thecorresponding limits established as a part of thetype certification of the engine and propellerinstalled in the airplane.

§ 3.745 Take-off operation. (a) Maximumrotational speed (revolutions per minute).

(b) Maximum permissible manifold pressure(if applicable).

(c) The time limit upon the use of thecorresponding power.

(d) Where the time limit of paragraph (c) ofthis section exceeds 2 minutes, the maximumallowable temperatures for cylinder head, oil, andcoolant outlet if applicable.

§ 3.746 Maximum continuous operation,(a) Maximum rotational speed (revolutions perminute).

(b) Maximum permissible manifold pressure(if applicable).

(c) Maximum allowable temperatures forcylinder head, oil, and coolant outlet if applicable.

§ 3.747 Fuel octane rating. The minimumoctane rating of fuel required for satisfactoryoperation of the power plant at the limits of §§3.745 and 3.746.

AIRPLANE WEIGHT

§ 3.748 Airplane weight. The airplaneweight and center of gravity limitations are thoserequired to be determined by § 3.71.

MINIMUM FLIGHT CREW

§ 3.749 Minimum flight crew. The minimumflight crew shall be established as that number ofpersons required for the safe operation of theairplane during any contact flight as determinedby the availability and satisfactory operation ofall necessary controls by each operatorconcerned.

TYPES OF OPERATION

§ 3.750 Types of operation. The type ofoperation to which the airplane is limited shall beestablished by the category in which it has beenfound eligible for certification and by theequipment installed. (See Parts 42 and 43 of thischapter.)

MARKINGS AND PLACARDS

§ 3.755 Markings and placards. (a) Themarkings and placards specified are required forall airplanes. Placards shall be displayed in aconspicuous place and both shall be such thatthey cannot be easily erased, disfigured, orobscured. Additional informational placards andinstrument markings having a direct andimportant bearing on safe operation may berequired by the Administrator when unusualdesign, operating, or handling characteristics sowarrant.

(b) When an airplane is certificated in morethan one category, the applicant shall select onecategory on which all placards and markings onthe airplane shall be based. The placard andmarking information for the other categories inwhich the airplane is certificated shall be enteredin the Airplane Flight Manual. A reference tothis information shall be included on a placardwhich shall also indicate the category on whichthe airplane placards and markings are based.

INSTRUMENT MARKINGS

§ 3.756 Instrument markings. Theinstruments listed in §§ 3.757-3.761 shall have thefollowing limitations marked thereon. Whenthese markings are placed on the cover glass ofthe instrument, adequate provision shall be madeto maintain the correct alignment of the glasscover with the face of the dial. All arcs and linesshall be of sufficient width and so located as tobe clearly and easily visible to the pilot.

§ 3.757 Air-speed indicator. (a) Trueindicated air speed shall be used:

(1) The never-exceed speed, Vne—a radialred line (see § 3.739).

(2) The caution range—a yellow arcextending from the red line in (1) above to theupper limit of the green arc specified in (3) below.

(3) The normal operating range—a greenarc with the lower limit at Vs1, as determined in §3.82 with maximum weight, landing gear and wingflaps retracted, and the upper limit at themaximum structural cruising speed established in§ 3.740.

(4) The flap operating range—a white arcwith the lower limit at Vso as determined in § 3.82at the maximum weight, and the upper limit at theflaps-extended speed in § 3.742.

(b) When the never-exceed and maximumstructural cruising speeds vary with altitude,means shall be provided which will indicate theappropriate limitations to the pilot throughout theoperating altitude range.

§ 3.758 Magnetic direction indicator. Aplacard shall be installed on or in close proximityto the magnetic direction indicator whichcontains the calibration of the instrument in alevel flight attitude with engine(s) operating andradio receiver(s) on or off (which shall be stated).The calibration readings shall be those to knownmagnetic headings in not greater than 30-degreeincrements.

§ 3.759 Power-plant instruments. Allrequired power-plant instruments shall be markedwith a red radial line at the maximum and minimum(if applicable) indications for safe operation. Thenormal operating ranges shall be marked with agreen arc which shall not extend beyond themaximum and minimum limits for continuousoperation. Take-off and precautionary rangesshall be marred with a yellow arc.

§ 3.760 Oil quantity indicators. Indicatorsshall be suitably marked in sufficient incrementsso that they will readily and accurately indicatethe quantity of oil.

§ 3.761 Fuel quantity indicator. When theunusable fuel supply for any tank exceeds 1gallon or 5 percent of the tank capacity,whichever is greater, a red band shall be placedon the indicator extending from the calibratedzero reading (see § 3.437) to the lowest readingobtainable in the level flight attitude, and asuitable notation in the Airplane Flight Manualshall be provided to indicate the flight personnelthat the fuel remaining in the tank when thequantity indicator reaches zero cannot be usedsafely in flight. (See § 3.672.)

CONTROL MARKINGS

§ 3.762 General. All cockpit controls, withthe exception of the primary flight controls, shallbe plainly marked as to their function and methodof operation.

§ 3.763 Aerodynamic controls. Thesecondary controls shall be suitably marked tocomply with §§ 3.337 and 3.338.

§ 3.764 Power-plant fuel controls. (a)Controls for fuel tank selector valves shall be

marked to indicate the position corresponding toeach tank and to all existing cross feed positions.

(b) When more than one fuel tank isprovided, and if safe operation depends upon theuse of tanks in a specific sequence, the fuel tankselector controls shall be marked adjacent to oron the control to indicate to the flight personnelthe order in which the tanks must be used.

(c) On multiengine airplanes, controls forengine valves shall be marked to indicate theposition corresponding to each engine.

(d) The capacity of each tank shall beindicated adjacent to or on the fuel tank selectorcontrol.

§ 3.765 Accessory and auxiliary controls.(a) When a retractable landing gear is used, theindicator required in § 3.359 shall be marked insuch a manner that the pilot can ascertain at alltimes when the wheels are secured in the extremepositions.

(b) Emergency controls shall be colored redand clearly marked as to their method ofoperation.

MISCELLANEOUS

§ 3.766 Baggage compartments, ballastlocation, and special seat loading limitations.(a) Each baggage or cargo compartment andballast location shall bear a placard which statesthe maximum allowable weight of contents and, ifapplicable, any special limitation of contents dueto loading requirements, etc.

(b) When the maximum permissible weightto be carried in a seat is less than 170 pounds(see § 3.74), a placard shall be permanentlyattached to the seat structure which states themaximum allowable weight of occupants to becarried.

§ 3.767 Fuel, oil, and coolant filleropenings. The following information shall bemarked on or adjacent to the filler cover in eachcase:

(a) The word "fuel," the minimumpermissible fuel octane number for the enginesinstalled, and the usable fuel tank capacity. (See§ 3.437.)

(b) The word "oil" and the oil tank capacity.

(c) The name of the proper coolant fluid

and the capacity of the coolant system.

§ 3.768 Emergency exit placards.Emergency exit placards and operating controlsshall be colored red. A placard shall be locatedadjacent to the control(s) which clearly indicatesit to be an emergency exit and describes themethod of operation. (See § 3.387.)

§ 3.769 Approved flight maneuvers—(a)Category N . A placard shall be provided in frontof and in clear view of the pilot stating: "Noacrobatic maneuvers including spins approved."

(b) Category U. A placard shall beprovided in front of and in clear view of the pilotstating: "No acrobatic maneuvers approved,except those listed in the Airplane FlightManual."

(c) Category A. A placard shall beprovided in clear view of the pilot which lists allapproved acrobatic maneuvers and therecommended entry air speed for each. Ifinverted flight maneuvers are not approved, theplacard shall bear a notation to this effect.

§ 3.770 Airplane category placard. Aplacard shall be provided in front of and in clearview of the pilot stating: "This airplane must beoperated as a ------------------ or ----------------category airplane in compliance with the AirplaneFlight Manual."

AIRPLANE FLIGHT MANUAL

§ 3.777 Airplane Flight Manual. AnAirplane Flight Manual shall be furnished witheach airplane. The portions of this documentlisted below shall be verified and approved bythe Administrator, and shall be segregated,identified, and clearly distinguished fromportions not so approved. Additional items ofinformation having a direct and important bearingon safe operation may be required by theAdministrator when unusual design, operating,or handling characteristics so warrant.

§ 3.778 Operating limitations—(a) Air-speed limitations. Sufficient information shall beincluded to permit proper marking of the air-speed limitations on the indicator as required in §3.757. It shall also include the design,maneuvering speed, and the maximum safe airspeed at which the landing gear can be safelylowered. In addition to the above information,the significance of the air speed limitations and of

the color coding used shall be explained.

(b) Power-plant limitations. Sufficientinformation shall be included to outline andexplain all power-plant limitations (see § 3.744)and to permit marking the instruments as requiredin § 3.759.

(c) Weight. The following information shallbe included:

(1) Maximum weight for which the airplanehas been certificated,

(2) Airplane empty weight and center ofgravity location,

(3) Useful load,

(4) The composition of the useful load,including the total weight of fuel and oil withtanks full.

(d) Load distribution. (1) All authorizedcenter of gravity limits shall be stated. If theavailable space for loading the airplane isadequately placarded or so arranged that anyreasonable distribution of the useful load listed inweight above will not result in a center of gravitylocation outside of the stated limits, this sectionneed not include any other information than thestatement of center of gravity limits.

(2) In all other cases this section shall alsoinclude adequate information to indicatesatisfactory loading combinations which willassure maintaining the center of gravity positionwithin approved limits.

(e) Maneuvers. All authorized maneuversand the appropriate air-speed limitations as wellas all unauthorized maneuvers shall be includedin accordance with the following:

(1) Normal category. All acrobaticmaneuvers, including spins, are unauthorized. Ifthe airplane has been demonstrated to becharacteristically incapable of spinning inaccordance with § 3.124 (d), a statement to thiseffect shall be entered here.

(2) Utility category. All authorizedmaneuvers demonstrated in the type flight testsshall be listed, together with recommended entryspeeds. All other maneuvers are not approved.If the airplane has been demonstrated to becharacteristically incapable of spinning inaccordance with § 3.124 (d), a statement to thiseffect shall be entered here.

(3) Acrobatic category. All approved flightmaneuvers demonstrated in the type flight testsshall be included, together with recommendedentry speeds.

(f) Flight load factor. The positive limitload factors made good by the airplane’sstructure shall be described here in terms ofaccelerations.

(g) Flight crew. When a flight crew of morethan one is required to operate the airplanesafely, the number and functions of this minimumflight crew shall be included.

§ 3.779 Operating procedures. Thissection shall contain information concerningnormal and emergency procedures and otherpertinent information peculiar to the airplane’soperating characteristics which are necessary tosafe operation.

§ 3.780 Performance information. (a)Information relative to the following items ofperformance shall be included:

(1) The stalling speed, Vso, at maximumweight,

(2) The stalling speed, Vs1, at maximumweight and with landing gear and wing flapsretracted,

(3) The take-off distance determined inaccordance with § 3.84, including the air speed atthe 50-foot height, and the airplane configuration,if pertinent,

(4) The landing distance determined inaccordance with § 3.86, including the airplaneconfiguration, if pertinent,

(5) The steady rate of climb determined inaccordance with § 3.85 (a), (c), and, asappropriate, (b), including the air speed, power,and airplane configuration, if pertinent.

(b) The effect of variation in (a) (2) withangle of bank up to 60 degrees shall be included.

(c) The calculated approximate effect ofvariations in subparagraphs (3), (4) and (5) of thisparagraph with altitude and temperature shall beincluded.

S UBPART H——IDENTIFICATION DATA

§ 3.791 Name plate. A name plate shall besecurely attached to and located in the pilotcompartment which shall contain:

(a) The manufacturer’s name and address.

(b) Model and serial numbers.

(c) Date of manufacture.

(d) Type certificate number.

(e) Production certificate number, (ifpertinent).

§ 3.792 Airworthiness certificate number.The identifying symbols and registrationnumbers shall be permanently affixed to theairplane structure in compliance with § 43.10 (c)of this chapter.

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