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FAR Final Rule
 Federal Register Information
[Federal Register: December 18, 1964]
[Page 17955]
Header Information
DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 23
[Docket No. 4080; Amendment No. 23-0]
Airworthiness Standards: Normal, Utility, and Acrobatic Category Airplanes [New]
Preamble Information
AGENCY: Federal Aviation Administration, DOT
ACTION: Final Rule
SUMMARY: This amendment adds Part 33 [New] to the Federal Aviation regulations to replace Part 3 of the Civil Air Regulations and is a part of the Agency recodification program announced in Draft Release 61-25, published in the Federal Register on November 15, 1961 (26 F.R. 10698).
EFFECTIVE DATE: This rule becomes effective February 1, 1965.
SUPPLEMENTARY INFORMATION:
Part 23 [New] was published as a notice of proposed rule making in the Federal Register on April 14, 1964 (29 F.R. 5111), and given further distribution as Notice No. 64-17.
During the life of the recodification project, Chapter I of Title 14 may contain more than one part bearing the same number. To differentiate between the two, the recodified parts, such as this one, will be labeled "[New]". The label will of course be dropped at the completion of the project as all of the regulations will be new.
Many of the comments received recommended specific substantive changes to the regulations. Although many of the recommendations appear to be meritorious, they cannot be adopted as a part of the recodification program. The purpose of the program is simply to streamline and clarify present regulatory language and delete obsolete or redundant provisions. To attempt substantive changes, other than relaxatory ones that are completely noncontroversial, would delay the project and be contrary to the ground rules specified for it in Draft Release 61-25. However, we recognize that an overall substantive review of the Part is long overdue. This review is now being undertaken and all substantive comments received are being carefully studied.
Present CAR Part 3 reflects the various writing styles used by those who have worked on it in the past. The recodification has allowed us to use one style throughout Part 23 [New]. The style changes that have been made do not affect substance. They have been made to ensure consistency in language throughout the new Federal Aviation Regulations, thereby making them easier to understand and apply. Part 23 [New] substitutes the word "must" for "shall". This has been done to reflect the fact that airworthiness standards are simply conditions precedent that are required to be met for the issue of a type certificate. The imperative "shall" would be inappropriate in this case. The failure to meet the standards simply results in a denial of the issue of the type certificate.
The sections in Part 23 [New] have been rearranged and renumbered. This will allow the requirements of this Part to be numbered identically with the comparable requirements in Parts 25 [New], 27 [New], and 29 [New]. In addition, some material has been rearranged in order that the requirements be more logically placed within the Part. An example of this is the regrouping of the requirements in Secs. 3.73, 3.73-3, 3.76, 3.84, 3.86, 3.120, 3.123, 3.124, and 3.437 through 3.780, dealing with recording of data and information, into the division of Part 23 [New] dealing with flight manual requirements. A similar rearrangement was the combining of the flutter requirements of Secs. 3.159 and 3.311 into Sec. 23.629.
As was stated in the preamble of the notice of proposed rule making of Part 23, those definitions in present Part 3 (and not now in Part 1 or executed in this part) that are necessary, will be recodified with the definitions of other airworthiness parts and added to Part 1 [New].
FAR 23 [New] contains, in addition to the CAM material included in the notice of proposed rule making, CAMs 3.71-1 and 3.311-1, and the second sentence of CAM 3.422-2. CAM 3.71-1 was included as it relaxes the rule by allowing certain tolerances during flight testing. These tolerances are necessary for the proper conduct of flight testing and such tolerances have been safely used in the past. They are now specifically included in the rule on flight testing. The flutter prevention method described in CAM 3.311-1 has been an additional safe acceptable method meeting the flutter requirements of CAR 3.311 and therefore has been included in this part. The second sentence of Sec. 3.422-2, with regard to the amount of deflection necessary to show propeller clearance for airplanes with leaf spring shock struts, has been included in order to eliminate the need for extensive testing ranges for such airplanes. The deflection corresponding to 1.5g has been found to be safe and reliable for this purpose and has therefore been included in the rule. That CAM material that has not been incorporated in Part 23 [New] has been determined to be advisory, not regulatory, in nature. This material is being reviewed, and where current and necessary, it will be issued in the Agency's Advisory Circular System.
In the table in present Sec. 3.106, yaw values are given for the "stick" and "wheel" when in fact the yaw values are applicable only to rudder pedal application. The table in this part has been revised accordingly.
The requirement in present CAM 8.294 with regard to approved bolts, pins, screws, and rivets and locking devices for them has been deleted as unnecessary as the Agency does not require specific approval of these items.
Paragraphs A23.7(e)(1) and (3) of Appendix A were reworded, in light of comments received, to make them consistent with the structural load criteria in the basic part.
Other minor changes of a technical clarifying nature have been made. They are not substantive and do not impose any burden on regulated persons.
The definitions, abbreviations, and rules of construction in Part 1 [New] of the Federal Aviation regulations apply to Part 23 [New].
Interested persons have been afforded an opportunity to participate in the making of this regulation and due consideration has been given to all relevant matter presented. The Agency is particularly appreciative of the cooperative spirit in which the public's comments were submitted.
Regulatory Information
In consideration of the foregoing, Chapter I of Title 14 is amended as follows, effective February 1, 1965:
1. By deleting Part 3.
2. By adding a Part 23 [New] reading as hereinafter set forth.
Subpart A--General
Sec.
23.1 Applicability.
23.3 Airplane categories.
Subpart B--Flight
General
23.21 Proof of compliance.
23.23 Load distribution limits.
23.25 Weight limits.
23.29 Empty weight and corresponding center of gravity.
23.31 Removable ballast.
23.33 Propeller speed and pitch limits.
Performance
23.45 General.
23.49 Stalling speed.
23.51 Takeoff.
23.65 Climb: all engines operating.
23.67 Climb: one engine inoperative
23.75 Landing.
23.77 Balked landing.
Flight Characteristics
23.141 General.
Controllability and Maneuverability
23.143 General.
23.145 Longitudinal control.
23.147 Directional and lateral control.
23.149 Minimum control speed.
23.151 Acrobatic maneuvers.
Trim
23.161 Trim.
Stability
23.171 General.
23.173 Static longitudinal stability.
23.175 Demonstration of static longitudinal stability.
23.177 Directional and lateral stability.
23.179 Instrumented stick force measurements.
23.181 Dynamic longitudinal stability.
Stalls
23.201 Stall demonstration.
23.203 Stall characteristics.
23.205 Stalls: critical engine inoperative.
23.207 Stall warning.
Spinning
23.221 Spinning.
Ground and Water Handling Characteristics
23.231 Longitudinal stability and control.
23.233 Directional stability and control.
23.235 Taxiing condition.
23.239 Spray characteristics.
Miscellaneous Flight Requirements
23.251 Vibration and buffeting.
Subpart C--Structure
General
23.301 Loads.
23.303 Factor of safety.
23.305 Strength and deformation.
23.307 Proof of structure.
Flight Loads
23.321 General.
23.331 Symmetrical flight conditions.
23.333 Flight envelope.
23.335 Design airspeeds.
23.337 Limit maneuvering load factors.
23.341 Gust load factors.
23.345 High lift devices.
23.347 Unsymmetrical flight conditions.
23.349 Rolling conditions.
23.351 Yawing conditions.
23.361 Engine torque.
23.363 Side load on engine mount.
23.365 Pressurized cabin loads.
23.369 Special conditions for rear lift truss.
Control Surface and System Loads
23.391 Control surface loads.
23.395 Control system.
23.397 Control system loads.
23.399 Dual control system.
23.405 Secondary control system.
23.407 Trim tab effects.
23.409 Tabs.
23.415 Ground gust conditions.
Horizontal Tail Surfaces
23.421 Balancing loads.
23.423 Maneuvering loads.
23.425 Gust loads.
23.427 Unsymmetrical loads.
Vertical Tail Surfaces
23.441 Maneuvering loads.
23.443 Gust loads.
23.445 Outboard fins.
Ailerons, Wing Flaps, and Special Devices
23.455 Ailerons.
23.457 Wing flaps.
23.459 Special devices.
Ground Loads.
23.471 General.
23.473 Ground load conditions and assumptions.
23.477 Landing gear arrangement.
23.479 Level landing conditions.
23.481 Tail down landing conditions.
23.483 One-wheel landing conditions.
23.485 Side load conditions.
23.493 Braked roll conditions.
23.497 Supplementary conditions for tail wheels.
23.499 Supplementary conditions for nose wheels.
23.505 Supplementary conditions for ski-planes.
Water Loads
23.521 Water load conditions.
Emergency Landing Conditions
23.561 General.
Fatigue Evaluation
23.571 Pressurized cabin.
Subpart D--Design and Construction
23.601 General.
23.603 Materials and workmanship.
23.605 Fabrication methods.
23.607 Self-locking nuts.
23.609 Protection of structure.
23.611 Inspection provisions.
23.613 Material strength properties and design values.
23.615 Design properties.
23.617 Interchangeability of seamwelded and seamless steel tubing.
23.619 Special factor.
23.621 Casting factors.
23.623 Bearing factors.
23.625 Fitting factors.
23.627 Fatigue strength.
23.629 Flutter.
Wings
23.641 Proof of strength.
23.643 Rib tests.
Control Surfaces
23.651 Proof of strength.
23.655 Installation.
23.657 Hinges.
23.659 Mass balance.
Control Systems
23.671 General.
23.673 Primary flight controls.
23.675 Stops.
23.677 Trim systems.
23.679 Control system locks.
23.681 Limit load static tests.
23.683 Operation tests.
23.685 Control system details.
23.687 Spring devices.
23.689 Cable systems.
23.693 Joints.
23.697 Wing flap controls.
23.699 Wing flap position indicator.
23.701 Flap interconnection.
Landing Gear
23.721 General.
23.723 Shock absorption tests.
23.725 Limit drop tests.
23.727 Reserve energy absorption drop test.
23.729 Retracting mechanism.
23.731 Wheels.
23.733 Tires.
23.735 Brakes.
23.737 Skis.
Floats and Hulls
23.751 Main float buoyancy.
23.753 Main float design.
23.755 Hulls.
23.757 Auxiliary floats.
Personnel and Cargo Accommodations
23.771 Pilot compartment.
23.773 Pilot compartment view.
23.775 Windshields and windows.
23.777 Cockpit controls.
23.779 Motion and effect of cockpit controls.
23.781 Cockpit control knob shape.
23.783 Doors.
23.785 Seats and berths.
23.787 Cargo compartments.
23.807 Emergency exits.
23.831 Ventilation.
Pressurization
23.841 Pressurized cabins.
23.843 Pressurization tests.
Fire Protection
23.853 Compartment interiors.
23.859 Combustion heater fire protection.
Miscellaneous
23.871 Leveling marks.
Subpart E--Powerplant
General
23.901 Installation.
23.903 Engines.
23.905 Propellers.
23.907 Propeller vibration.
23.925 Propeller clearance.
Fuel System
23.951 General.
23.953 Fuel system independence.
23.955 Fuel flow.
23.957 Flow between interconnected tanks.
23.959 Unusable fuel supply and fuel system operation on low fuel.
23.961 Fuel system hot weather operation.
23.963 Fuel tanks: general.
23.965 Fuel tank tests.
23.967 Fuel tank installation.
23.969 Fuel tank expansion space.
23.971 Fuel tank sump.
23.973 Fuel tank filler connection.
23.975 Fuel tank vents and carburetor vapor vents.
23.977 Fuel tank outlet.
Fuel System Components
23.991 Fuel pumps.
23.993 Fuel system lines and fittings.
23.995 Fuel valves and controls.
23.997 Fuel strainer or filter.
23.999 Fuel system drains.
Oil System
23.1011 General.
23.1013 Oil tanks.
23.1015 Oil tank tests.
23.1017 Oil lines and fittings.
23.1019 Oil strainer or filter.
23.1021 Oil system drains.
23.1023 Oil radiators.
23.1027 Propeller feathering system.
Cooling
23.1041 General.
23.1043 Cooling tests.
23.1045 Cooling test procedures for single-engine airplanes.
23.1047 Cooling test procedures for multi-engine airplanes.
Liquid Cooling
23.1061 Installation.
23.1063 Coolant tank tests.
Induction System
23.1091 Air induction.
23.1093 Induction system icing protection.
23.1095 Carburetor deicing fluid flow rate.
23.1097 Carburetor deicing fluid system capacity.
23.1099 Carburetor deicing fluid system detail design.
23.1101 Carburetor air preheater design.
23.1103 Induction system ducts.
23.1105 Induction system screens.
Exhaust System
23.1121 General.
23.1123 Exhaust manifold.
23.1125 Exhaust heat exchangers.
Powerplant Controls and Accessories
23.1141 Powerplant controls: general.
23.1143 Throttle controls.
23.1145 Ignition switches.
23.1147 Mixture controls.
23.1149 Propeller speed and pitch controls.
23.1153 Propeller feathering controls.
23.1157 Carburetor air temperature controls.
23.1163 Powerplant accessories.
23.1165 Engine ignition systems.
Powerplant Fire Protection
23.1183 Lines and fittings.
23.1189 Shutoff means.
23.1191 Firewalls.
23.1193 Cowling.
Subpart F--Equipment
General
23.1301 Function and installation.
23.1303 Flight and navigation instruments.
23.1305 Powerplant instruments.
23.1307 Miscellaneous equipment.
Instrument: Installation
23.1321 Arrangement and visibility.
23.1323 Airspeed indicating system.
23.1325 Static air vent system.
23.1327 Magnetic direction indicator.
23.1329 Automatic pilot system.
23.1331 Instruments using a power supply.
23.1335 Flight director instrument.
23.1337 Powerplant instruments.
Electrical Systems and Equipment
23.1351 General.
23.1353 Storage battery design and installation.
23.1357 Circuit protective devices.
23.1361 Master Switch.
23.1365 Electric cables.
23.1367 Switches.
Lights
23.1381 Instrument lights.
23.1383 Landing lights.
23.1385 Position light system installation.
23.1387 Position light system dihedral angles.
23.1389 Position light distribution and intensities.
23.1391 Minimum intensities in the horizontal plane of forward and rear position lights.
23.1393 Minimum intensities in any vertical plane of forward and rear position lights.
23.1395 Maximum intensities in overlapping beams of forward and rear position lights.
23.1397 Color specifications.
23.1399 Riding light.
23.1401 Anticollision light system.
Safety Equipment
23.1411 Accessibility.
23.1413 Safety belts.
23.1415 Ditching equipment.
23.1419 Deicers.
Miscellaneous Equipment
23.1431 Electronic equipment.
23.1435 Hydraulic systems.
23.1437 Accessories for multiengine airplanes.
Subpart G--Operating Limitations and information
23.1501 General.
23.1505 Airspeed limitations.
23.1507 Maneuvering speed.
23.1511 Flap extended speed.
23.1513 Minimum control speed.
23.1519 Weight and center of gravity.
23.1521 Powerplant limitations.
23.1523 Minimum flight crew.
23.1525 Kinds of operation.
Markings and Placards
23.1541 General.
23.1543 Instrument markings: general.
23.1545 Airspeed indicator.
23.1547 Magnetic direction indicator.
23.1549 Powerplant instruments.
23.1551 Oil quantity indicator.
23.1553 Fuel quantity indicator.
23.1555 Control markings.
23.1557 Miscellaneous markings and placards.
23.1559 Operating limitations placard.
23 1561 Safety equipment.
23.1563 Airspeed placards
23.1567 Flight maneuver placard.
Airplane Flight Manual
23.1581 General.
23.1583 Operating limitations.
23.1585 Operating procedures.
23.1587 Performance information.
23.1589 Loading information.
Appendix A--Simplified Design Load Criteria for Conventional, Single-Engine Airplanes of 6,000 Pounds or Less Maximum Weight
Sec.
A23.1 General.
A23.3 Special symbols.
A23.5 Certification in more than one category.
A23.7 Flight loads.
A23.9 Flight conditions.
A23.11 Control surface loads.
A23.13 Control system loads.
Appendix B--Control Surface Loadings
B23.1 General.
B23.11 Control surface loads.
Appendix C--Basic Landing Conditions
C23.1 Basic landing conditions.
Appendix D--Wheel Spin-Up Loads
D23.1 Wheel spin-up loads.
Authority: The provisions of this Part 23 issued under Secs. 313(a), 601, 603, Federal Aviation Act of 1958; 49 U.S.C. 1354(a), 1421, 1423.
Subpart A--General
Sec. 23.1 Applicability.
(a) This part prescribes airworthiness standards for the issue of type certificates, and changes to those certificates, for small airplanes in the normal, utility, and acrobatic categories.
(b) Each person who applies under Part 21 [New] for such a certificate or change must show compliance with the applicable requirements of this part.
Sec. 23.3 Airplane categories.
(a) The normal category is limited to airplanes intended for nonacrobatic operation. Nonacrobatic operation includes any maneuvers incident to normal flying, stalls (except whip stalls), and turns in which the angle of bank is not more than 60 degrees.
(b) The utility category is limited to airplanes intended for limited acrobatic operation. Limited acrobatic operation includes any maneuvers incident to normal flying, stalls (except whip stalls), spins (if approved for the particular type of airplane), lazy eights, chandelles, and steep turns in which the angle of bank is more than 60 degrees.
(c) The acrobatic category is limited to airplanes intended for use without restrictions other than those shown to be necessary as a result of required flight tests.
(d) Small airplanes may be certificated in more than one category if the requirements of each requested category are met.
Subpart B--Flight
Sec. 23.21 Proof of compliance.
(a) Each requirement of this subpart must be met at each appropriate combination of weight and center of gravity within the range of loading conditions for which certification is requested. This must be shown--
(1) By tests upon an airplane of the type for which certification is requested, or by calculations based on, and equal in accuracy to, the results of testing; and
(2) By systematic investigation of each probable combination of weight and center of gravity, if compliance cannot be reasonably inferred from combinations investigated.
(b) The following general tolerances are allowed during flight testing. However, greater tolerances may be allowed in particular tests:
| Item | Tolerance |
| Weight------------------------------------------------------ | +5%, -10% |
| Critical items affected by weight------------------ | +5%, -1% |
| C.G.---------------------------------------------------------- | ±7% total travel. |
Sec. 23.23 Load distribution limits.
Ranges of weights and centers of gravity within which the airplane may be safely operated must be established. If low fuel adversely affects balance or stability, the airplane must be tested under conditions simulating those that would exist when the amount of usable fuel does not exceed one gallon for each 12 maximum continuous horsepower of the engine or engines.
Sec. 23.25 Weight limits.
(a) Maximum weight. The maximum weight (the highest weight at which compliance with each applicable requirement of this part is shown) must be established so that it is--
(1) Not more than--
(i) The highest weight selected by the applicant;
(ii) Except as provided in Sec. 23.473 for multiengine airplanes, the design maximum weight (the highest weight at which compliance with each applicable structural loading condition of this part is shown); or
(iii) The highest weight at which compliance with each applicable flight requirement of this part is shown; and
(2) Assuming a weight of 170 pounds for each occupant of each seat for normal category airplanes and 190 pounds (unless otherwise placarded) for utility and acrobatic category airplanes, not less than the weight with--
(i) Each seat occupied, oil at full tank capacity, and at least enough fuel for one-half hour of operation at rated maximum continuous power; or
(ii) The required minimum crew, and fuel and oil to full tank capacity.
(b) Minimum weight. The minimum weight (the lowest weight at which compliance with each applicable requirement of this part is shown) must be established so that it is not more than the sum of--
(1) The empty weight determined under Sec. 23.29;
(2) The weight of the required minimum crew (assuming a weight of 170 pounds for each crewmember);
(3) The weight of the oil at full tank capacity; and
(4) The weight of no more than the quantity of fuel necessary for one-half hour of operation at rated maximum continuous power.
Sec. 23.29 Empty weight and corresponding center of gravity.
(a) The empty weight and corresponding center of gravity must be determined by weighing the airplane with--
(1) Fixed ballast;
(2) Unusable fuel determined under Sec. 23.959;
(3) Undrainable oil (the oil remaining in the airplane while in the ground attitude after drainage of all drainable oil in that attitude);
(4) Engine coolant; and
(5) Hydraulic fluid.
(b) The condition of the airplane at the time of determining empty weight must be one that is well defined and can be easily repeated.
Sec. 23.31 Removable ballast.
Removable ballast may be used in showing compliance with the flight requirements of this subpart, if--
(a) The place for carrying ballast is properly designed and installed, and is marked under Sec. 23.1557; and
(b) The Airplane Flight Manual includes instructions for the proper placement of the removable ballast under each loading condition for which removable ballast is necessary.
Sec. 23.33 Propeller speed and pitch limits.
(a) General. The propeller speed and pitch must be limited to values that will assure safe operation under normal operating conditions.
(b) Propellers not controllable in flight. For each propeller whose pitch cannot be controlled in flight--
(1) During takeoff and initial climb at VY, the propeller must limit the engine r.p.m., at full throttle or at maximum allowable takeoff manifold pressure, to a speed not greater than the maximum allowable takeoff r.p.m.; and
(2) During a closed throttle glide at the placarded "never-exceed speed", the propeller may not cause an engine speed above 110 percent of maximum continuous speed.
(c) Controllable pitch propellers without constant speed controls. Each propeller that can be controlled in flight, but that does not have constant speed controls, must have a means to limit the pitch range so that--
(1) The lowest possible pitch allows compliance with paragraph (b)(1) of this section; and
(2) The highest possible pitch allows compliance with paragraph (b)(2) of this section.
(d) Controllable pitch propellers with constant speed controls. Each controllable pitch propeller with constant speed controls must have--
(1) With the governor in operation, a means at the governor to limit the maximum engine speed to the maximum allowable takeoff r.p.m.; and
(2) With the governor inoperative, a means to limit the maximum engine speed to 103 percent of the maximum allowable takeoff r.p.m. with the propeller blades at the lowest possible pitch and with takeoff manifold pressure, the airplane stationary, and no wind.
Performance
Sec. 23.45 General.
Compliance with the performance requirements of this subpart must be shown for still air with a standard atmosphere.
Sec. 23.49 Stalling speed.
(a)  is the stalling speed, if obtainable, or the minimum steady speed, in miles per hours (CAS), at which the airplane is controllable, with the--
(1) Engines idling, throttles closed (or at not more than the power necessary for zero thrust at a speed not more than 110 percent of the stalling speed);
(2) Propellers in the takeoff position;
(3) Landing gear extended;
(4) Wing flaps in the landing position;
(5) Cowl flaps closed;
(6) Center of gravity in the most unfavorable position within the allowable landing range; and
(7) Weight used when  is being used as a factor to determine compliance with a required performance standard.
(b)  at maximum weight may not exceed 70 miles per hour for--
(1) Single-engine airplanes; and
(2) Multiengine airplanes of 6,000 pounds or less maximum weight that cannot meet the minimum rate of climb specified in Sec. 23.67(b) with the critical engine inoperative.
(c)  is the calibrated stalling speed, if obtainable, or the minimum steady speed, in miles per hour, at which the airplane is controllable, with the--
(1) Engines idling, throttles closed (or at not more than the power necessary for zero thrust at a speed not more than 110 percent of the stalling speed);
(2) Propellers in the takeoff position;
(3) Airplane in the condition existing in the test in which  is being used; and
(4) Weight used when  is being used as a factor to determine compliance with a required performance standard.
(d)  and  must be determined by flight tests, using the procedure specified in Sec. 23.201.
Sec. 23.51 Takeoff.
(a) For airplanes of more than 6,000 pounds maximum weight (except skiplanes for which landplane takeoff data has been determined under this paragraph and furnished in the Airplane Flight Manual)--
(1) The distance required to take off and climb over a 50-foot obstacle must be determined with--
(i) The engines operating within approved operating limitations; and
(ii) The cowl flaps in the normal takeoff position;
(2) Upon reaching a height of 50 feet above the takeoff surface level, the airplane must have reached a speed of not less than--
(i) 1.3  ; or
(ii) Any lesser speed, not less than VX plus 5 miles per hour, that is shown to be safe under any condition, including turbulence and complete engine failure;
(3) The starting point for measuring seaplane and amphibian takeoff distance may be the point at which a speed of not more than three miles per hour is reached; and
(4) No takeoff made to determine the data required by this section may require exceptional piloting skill or exceptionally favorable conditions.
(b) For airplanes of 6,000 pounds or less maximum weight--
(1) The takeoff may not require exceptional piloting skill;
(2) With takeoff power, there must be enough elevator control--
(i) For a tail-wheel type airplane, to maintain, at 0.8  , an attitude that will allow holding the airplane on the runway until a safe takeoff speed is reached; and
(ii) For a nose-wheel type airplane to raise the nose-wheel clear of the takeoff surface at 0.85  .
Sec. 23.65 Climb: all engines operating.
(a) For airplanes of more than 6,000 pounds maximum weight--
(1) Each airplane must have a steady rate of climb at sea level of at least 300 feet per minute and a steady angle of climb of at least 1:12 for land planes or 1:15 for seaplanes and amphibians with--
(i) Not more than maximum continuous power on each engine;
(ii) The landing gear retracted;
(iii) The wing flaps in the takeoff position; and
(iv) The cowl flaps in the position used in the cooling tests required by Secs 23.1041 through 23.1047;
(2) Each airplane with engines for which the takeoff and maximum continuous power ratings are identical and that has fixed-pitch, two-position, or similar propellers, may use a lower propeller pitch setting than that allowed by Sec. 23.33 to obtain rated engine r.p.m. at VX, if--
(i) The airplane shows marginal performance (such as when it can meet the rate of climb requirements of paragraph (a)(1) of this section but has difficulty in meeting the angle of climb requirements of paragraph (a)(1) of this section or of Sec. 23.77); and
(ii) Acceptable engine cooling is shown at the lower speed associated with the best angle of climb.
(b) Each airplane of 6,000 pounds or less maximum weight must have a steady rate of climb at sea level of at least 300 feet per minute, or 10 VS1 (that is, the number of feet per minute is obtained by multiplying the number of miles per hour by 10), whichever is greater, with--
(1) Takeoff power;
(2) The landing gear extended;
(3) The wing flaps in the takeoff position; and
(4) The cowl flaps in the position used in the cooling tests required by Secs. 23.1041 through 23.1047.
Sec. 23.67 Climb: one engine inoperative.
(a) Each multiengine airplane of more than 6,000 pounds maximum weight must be able to maintain a steady rate of climb of at least 0.02 VS02 (that is, the number of feet per minute is obtained by multiplying the square of the number of miles per hour by 0.02) at an altitude of 5,000 feet with the--
(1) Critical engine inoperative, and its propeller in the minimum drag position;
(2) Remaining engines at not more than maximum continuous power;
(3) Landing gear retracted;
(4) Wing flaps in the most favorable position; and
(5) Cowl flaps in the position used in the cooling tests required by Secs. 23.1041 through 23.1047.
(b) For multiengine airplanes of 6,000 pounds or less maximum weight, the following apply:
(1) Each airplane with a VS0 of more than 70 miles per hour must be able to maintain a steady rate of climb of at least 0.02 VS02 (that is, the number of feet per minute is obtained by multiplying the square of the number of miles per hour by 0.02), at an altitude of 5,000 feet with the--
(i) Critical engine inoperative and its propeller in the minimum drag position;
(ii) Remaining engines at not more than maximum continuous power;
(iii) Landing gear retracted;
(iv) Wing flaps in the most favorable position; and
(v) Cowl flaps in the position used in the cooling tests required by Secs. 23.1041 through 23.1047.
(2) For each airplane with a stalling speed of 70 miles per hour or less, the steady rate of climb at 5,000 feet must be determined with the--
(i) Critical engine inoperative and its propeller in the minimum drag position;
(ii) Remaining engines at not more than maximum continuous power;
(iii) Landing gear retracted;
(iv) Wing flaps in the most favorable position; and
(v) Cowl flaps in the position used in the cooling tests required by Secs. 23.1041 through 23.1047.
Sec. 23.75 Landing.
(a) For airplanes of more than 6,000 pounds maximum weight (except skiplanes for which landplane landing data have been determined under this paragraph and furnished in the Airplane Flight Manual), the horizontal distance required to land and come to a complete stop (or to a speed of approximately three miles per hour for seaplanes and amphibians) from a point 50 feet above the landing surface must be determined as follows:
(1) A steady gliding approach with a calibrated airspeed of at least 1.5  must be maintained down to the 50 foot height.
(2) The landing may not require exceptional piloting skill or exceptionally favorable conditions.
(3) The landing must be made without excessive vertical acceleration or tendency to bounce, nose over, ground loop, porpoise, or water loop.
(b) Airplanes of 6,000 pounds or less maximum weight must be able to be landed safely and come to a stop without exceptional piloting skill and without excessive vertical acceleration or tendency to bounce, nose over, ground loop, porpoise, or water loop.
Sec. 23.77 Balked landing.
For balked landings, each airplane with a maximum weight of--
(a) More than 6,000 pounds, must be able to maintain a steady angle of climb at sea level of at least 1:30 with--
(1) Takeoff power on each engine;
(2) The landing gear extended; and
(3) The wing flaps in the landing position, except that, if the flaps may safely be retracted in two seconds or less without loss of altitude and without sudden changes of angle of attack or exceptional piloting skill, they may be retracted; and
(b) 6,000 pounds or less, must be able to maintain a steady rate of climb at sea level of at least 200 feet per minute, or 5  (that is, the number of feet per minute is obtained by multiplying the number of miles per hour by five), whichever is greater, with--
(1) Takeoff power on each engine;
(2) The landing gear extended; and
(3) The wing flaps in the landing position, except that, if rapid retraction is possible with safety, without loss of altitude, and without sudden changes of angle of attack or exceptional piloting skill, they may be retracted.
Flight Characteristics
Sec. 23.141 General.
The airplane must meet the requirements of Secs. 23.143 through 23.221--
(a) At the normally expected operating altitudes;
(b) Under any critical loading conditions within the center of gravity range; and
(c) Unless otherwise specified, at the highest weight for which certification is requested.
Controllability and Maneuverability
Sec. 23.143 General.
(a) The airplane must be safely controllable and maneuverable during --
(1) Takeoff;
(2) Climb;
(3) Level flight;
(4) Dive; and
(5) Landing (power on and power off).
(b) It must be possible to make a smooth transition from one flight condition to another (including turns and slips) without exceptional piloting skill, alertness, or strength, and without danger of exceeding the limit load factor, under any probable operating condition (including, for multiengine airplanes, those conditions normally encountered in the sudden failure of any engine).
(c) If marginal conditions exist with regard to required pilot strength, the "strength of pilots" limits must be shown by quantitative tests. In no case may the limits exceed those prescribed in the following table:
Values in pounds of force as applied to the control wheel or rudder pedals | Pitch | Roll | Yaw |
| (a) For temporary application: |  | | |
| Stick--------------------------------------------- | 60 | 30 | -------------------------- |
| Wheel (applied to rim)------------------- | 75 | 60 | -------------------------- |
| Rudder Pedal------------------------------- | -------------------------- | -------------------------- | 150 |
| (b) For prolonged application. | 10 | 5 | 20 |
Sec. 23.145 Longitudinal control.
(a) It must be possible, at speeds below the trim speed, to pitch the nose downward so that
the rate of increase in airspeed allows prompt acceleration to the trim speed with--
(1) Maximum continuous power on each engine and the airplane trimmed at VX;
(2) Power off and the airplane trimmed at 1.5  or at the minimum trim speed, whichever is higher; and
(3) Wing flaps and landing gear (i) retracted, and (ii) extended.
(b) With the landing gear extended, no change in trim or exertion of more control force than can be readily applied with one hand for a short period of time may be required for the following maneuvers:
(1) With power off, flaps retracted, and the airplane trimmed at 1.5  , or at the minimum trim speed, whichever is higher, extend the flaps as rapidly as possible while maintaining the airspeed at approximately 40 percent above the instantaneous value of the stalling speed.
(2) Repeat subparagraph (1) of this paragraph except initially extend the flaps and then retract them as rapidly as possible.
(3) Repeat subparagraph (2) of this paragraph except with maximum continuous power.
(4) With power off, flaps retracted, and the airplane trimmed at 1.5  , or at the minimum trim speed, whichever is higher, apply takeoff power rapidly while maintaining the same airspeed.
(5) Repeat subparagraph (4) of this paragraph, except with the flaps extended.
(6) With power off, flaps extended, and the airplane trimmed at 1.5  , or at the minimum trim speed, whichever is higher, obtain and maintain airspeeds between 1.1  and either 1.7  or VF, whichever is lower.
(c) It must be possible, without exceptional piloting skill, to maintain approximately level flight when flap retraction from any position is made during steady horizontal flight at 1.1  with simultaneous application of not more than maximum continuous power.
(d) It must be possible, with a pilot control force of not more than 10 pounds, to maintain a speed of not more than 1.5  during a power-off glide with landing gear and wing flaps extended, and with--
(1) The most forward center of gravity approved for the maximum weight; and
(2) The most forward center of gravity approved for any weight.
(e) It must be possible, by using the normal flight and power controls except the primary longitudinal control, to control the descent of the airplane to a zero rate of descent and to an attitude suitable for a controlled landing, without exceptional piloting skill, alertness, or strength, and without exceeding the operational and structural limitations of the airplane.
Sec. 23.147 Directional and lateral control.
(a) For each multiengine airplane, it must be possible to make turns with 15 degrees of bank both towards and away from an inoperative engine, from a steady climb at 1.4  or VY with--
(1) One engine inoperative and its propeller in the minimum drag position;
(2) The remaining engines at not more than maximum continuous power;
(3) The rearmost allowable center of gravity;
(4) The landing gear (i) retracted, and (ii) extended;
(5) The flaps in the most favorable climb position; and
(6) Maximum weight.
(b) For each multiengine airplane, it must be possible, while holding the wings level within five degrees, to make sudden changes in heading safely in both directions. This must be shown at 1.4  or VY with heading changes up to 15 degrees (except that the heading change at which the rudder force corresponds to the limits specified in Sec. 23.143 need not be exceeded), with the--
(1) Critical engine inoperative and its propeller in the minimum drag position;
(2) Remaining engines at maximum continuous power;
(3) Landing gear (i) retracted, and (ii) extended;
(4) Flaps in the most favorable climb position; and
(5) Center of gravity at its rearmost allowable position.
Sec. 23.149 Minimum control speed.
(a) VMC is the minimum calibrated airspeed at which, when any engine is suddenly made inoperative, it is possible to recover control of the airplane with that engine still inoperative and maintain straight flight, either with zero yaw, or, at the option of the applicant, with an angle of bank of not more than five degrees. VMC may not exceed 1.2  with--
(1) Takeoff or maximum available power on each engine;
(2) The rearmost allowable center of gravity;
(3) The flaps in the takeoff position; and
(4) The landing gear retracted.
(b) At VMC, the rudder forces required to maintain control may not exceed the limitations set forth in Sec. 23.143, and it may not be necessary to throttle the remaining engines. During recovery, the airplane may not assume any dangerous attitude or require exceptional piloting skill, alertness, or strength, to prevent a heading change of more than 20 degrees.
Sec. 23.151 Acrobatic maneuvers.
Each acrobatic and utility category airplane must be able to perform safely the acrobatic maneuvers for which certification is requested. Safe entry speeds for these maneuvers must be determined.
Sec. 23.161 Trim.
(a) General. Each airplane must meet the trim requirements of this section after being trimmed, and without further pressure upon, or movement of, the primary controls or their corresponding trim controls by the pilot or the automatic pilot.
(b) Lateral and directional trim. The airplane must maintain lateral and directional trim in level flight at 0.9 VH or VC, whichever is lower, with the landing gear and wing flaps retracted.
(c) Longitudinal trim. The airplane must maintain longitudinal trim during--
(1) A climb with maximum continuous power at a speed between VX and 1.4  , with the landing gear and wing flaps retracted;
(2) A climb with maximum continuous power at a speed between VX and 1.4  , with the landing gear retracted and the wing flaps in the takeoff position;
(3) A power approach at 1.5  , with a three degree angle of descent, the landing gear extended, and the wing flaps retracted;
(4) A power approach at 1.5  , with a three degree angle of descent, the landing gear and wing flaps extended, and the most forward center of gravity approved for the maximum weight;
(5) A power approach at 1.5  , with a three degree angle of descent, the landing gear and wing flaps extended, and the most forward center of gravity approved for any weight; and
(6) Level flight at any speed from 0.9 VH to either VX or 1.4  , with landing gear and wing flaps retracted.
(d) In addition, each multiengine airplane must maintain longitudinal and directional trim at a speed between VY and 1.4  , with--
(1) The critical engine inoperative;
(2) The remaining engines at maximum continuous power;
(3) The landing gear retracted;
(4) Wing flaps retracted; and
(5) An angle of bank of not more than five degrees.
Stability
Sec. 23.171 General.
The airplane must be longitudinally, directionally, and laterally stable under Secs. 23.173 through 23.181. In addition, the airplane must show suitable stability and control "feel" (static stability) in any condition normally encountered in service, if flight tests show it is necessary for safe operation.
Sec. 23.173 Static longitudinal stability.
Under the conditions specified in Sec. 23.175 and with the airplane trimmed as indicated, the characteristics of the elevator control forces and the friction within the control system must be as follows:
(a) A pull must be required to obtain and maintain speeds below the specified trim speed and a push required to obtain and maintain speeds above the specified trim speed. This must be shown at any speed that can be obtained without excessive control force, except speeds more than the appropriate maximum allowable speed or less than the minimum speed for steady unstalled flight.
(b) The airspeed must return to within plus or minus 10 percent of the original trim speed when the control force is slowly released at any speed within the speed range specified in paragraph (a) of this section.
(c) The stick force must vary with speed so that any substantial speed change results in a stick force clearly perceptible to the pilot.
Sec. 23.175 Demonstration of static longitudinal stability.
Static longitudinal stability must be shown as follows:
(a) Climb. The stick force curve must have a stable slope at speeds between 1.2 4  , and 1.6 4  , with--
(1) Flaps retracted;
(2) Landing gear retracted;
(3) Maximum weight;
(4) 75 percent of maximum continuous power; and
(5) The airplane trimmed at 1.4 4  .
(b) Cruise. The stick force curve must have a stable slope at any speed obtainable with a stick force not more than 40 pounds at speeds between 1.3 4  and the maximum allowable speed, with--
(1) Landing gear retracted;
(2) Flaps retracted;
(3) Maximum weight;
(4) 75 percent of maximum continuous power; and
(5) The airplane trimmed for level flight.
Compliance with this paragraph must also be shown with the landing gear extended and without exceeding level flight trim speed.
(c) Approach. The stick force curve must have a stable slope and the stick force may not exceed 40 pounds at speeds between 1.1 4  and 1.8 4  with--
(1) Flaps in the landing position;
(2) Landing gear extended;
(3) Maximum weight; and
(4) The airplane trimmed at 1.5 4  with enough power to maintain a three degree angle of descent.
Sec. 23.177 Directional and lateral stability.
(a) Three-control airplanes. The stability requirements for three-control airplanes are as follows:
(1) The static directional stability, as shown by the tendency to recover from a skid with the rudder free, must be positive for any landing gear and flap position appropriate to the takeoff, climb, cruise, and approach configurations. This must be shown with symmetrical power up to maximum continuous power, and at speeds from 1.2  up to the maximum allowable speed for the condition being investigated. The angle of skid for these tests must be appropriate to the type of airplane. At larger angles of skid up to that at which full rudder is used or a control force limit in Sec. 23.143 is reached, whichever occurs first, and at speeds from 1.2  to VA, the rudder pedal force must not reverse.
(2) The static lateral stability, as shown by the tendency to raise the low wing in a slip, must be positive for any landing gear and flap positions. This must be shown with symmetrical power up to 75 percent of maximum continuous power at speeds above 1.2  , up to the maximum allowable speed for the configuration being investigated. The static lateral stability may not be negative at 1.2  . The angle of slip for these tests must be appropriate to the type of airplane, but in no case may the slip angle be less than that obtainable with 10 degrees of bank.
(3) In straight, steady slips at 1.2  for any landing gear and flap positions, and for any symmetrical power conditions up to 50 percent of maximum continuous power, the aileron and rudder control movements and forces must increase steadily (but not necessarily in constant proportion) as the angle of slip is increased up to the maximum appropriate to the type of airplane. At larger slip angles up to the angle at which the full rudder or aileron control is used or a control force limit contained in Sec. 23.143 is obtained, the rudder pedal force may not reverse. Enough bank must accompany slipping to hold a constant heading. Rapid entry into, or recovery from, a maximum slip may not result in uncontrollable flight characteristics.
(4) Any short period oscillation, occurring between stalling speed and the maximum allowable speed, must be heavily damped with the primary controls (i) free and (ii) in a fixed position.
(b) Two-control (or simplified control) airplanes. The stability requirements for two-control airplanes are as follows:
(1) The directional stability of the airplane must be shown by showing that, in each configuration, it can be rapidly rolled from a 45 degree bank in one direction to a 45 degree bank in the opposite direction without showing dangerous skid characteristics.
(2) The lateral stability of the airplane must be shown by showing that it will not assume a dangerous attitude or speed when the controls are abandoned for two minutes. This must be done in moderately smooth air with the airplane trimmed for straight level flight at 0.9 VH or VC, whichever is lower, with flaps and landing gear retracted, and with a rearward center of gravity.
(3) Any short period oscillation occurring between the stalling speed and the maximum allowable speed must be heavily damped with the primary controls (i) free and (ii) in a fixed position.
Sec. 23.179 Instrumented stick force measurements.
Instrumented stick force measurements must be made unless--
(a) Changes in speed are clearly reflected by changes in stick forces; and
(b) The maximum forces obtained under Secs. 23.173 and 23.175 are not excessive.
Sec. 23.181 Dynamic longitudinal stability.
Any short period longitudinal oscillation occurring between the stalling speed and the maximum allowable speed must be heavily damped with the primary controls (a) free and (b) fixed.
Stalls
Sec. 23.201 Stall demonstration.
(a) Level wing stalls must be shown with--
(1) Power off; and
(2) A power setting not less than that required to show compliance with Sec. 23.65 for an airplane of more than 6,000 lbs. maximum weight, or with 90 percent of maximum continuous power for an airplane of 6,000 lbs. or less maximum weight.
(b) In either condition required by paragraph (a) of this section, it must be possible to comply with the applicable requirements of Sec. 23.203 (a) with flaps and landing gear in any position.
(c) The following procedure must be used to show compliance with Sec. 23.203 (a):
(1) With the trim controls adjusted for straight flight at 1.5  , or at the minimum trim speed, whichever is higher, reduce the speed with the elevator control until speed is slightly above the stalling speed.
(2) Then pull back the elevator control so that the rate of speed reduction will not exceed one mile per hour per second until a stall is produced, as shown by an uncontrollable downward pitching motion of the airplane, or until the control reaches the stop. Normal use of the elevator control for recovery is allowed after the pitching motion has unmistakably developed.
(d) Except where made inapplicable by the special features of a particular type of airplane, the following procedure must be used to measure loss of altitude during a stall:
(1) The approach to the stall must be made as prescribed in paragraph (b) of this section.
(2) The loss of altitude encountered in the stall (power on or power off) is the change in altitude (as observed on the sensitive altimeter testing installation) between the altitude at which the airplane pitches and the altitude at which horizontal flight is regained.
(3) If required, the power used during stall recovery must be that which would be used under normal operating conditions in this maneuver. However, the power used to regain level flight may not be applied until flying control is regained.
(d) For turning flight stalls, the following maneuver must be used to show compliance with Sec. 23.203(b):
(1) Establish a steady, curvilinear, level, coordinated turn in a 30 degree bank and, while maintaining the 30 degree bank, stall the airplane by steadily and progressively tightening the turn with the elevator control until the airplane is stalled, or until the elevator has reached its stop.
(2) When the stall has fully developed, regain level flight with normal use of the controls.
Sec. 23.203 Stall characteristics.
(a) For level wing stalls--
(1) For an airplane with independently controlled rolling and directional controls, it must be possible to produce and to correct roll by unreversed use of the rolling control and to produce and correct yaw by unreversed use of the directional control, up to the time the airplane pitches in the maneuver prescribed in Sec. 23.201(b);
(2) For an airplane with interconnected lateral and directional controls (two control), for an airplane with only one of these controls, it must be possible to produce and correct roll by unreversed use of the rolling control without producing excessive yaw, up to the time the airplane pitches in the maneuver prescribed in Sec. 23.201(b); and
(3) During the recovery part of the maneuver prescribed in Sec. 23.201(b), it must be possible to prevent more than 15 degrees of roll or yaw by the normal use of the controls.
(b) For turning flight stalls, when stalled during a coordinated turn with 30 degrees of bank, 75 percent maximum continuous power on each engine, and flaps and landing gear retracted, it must be possible to regain normal level flight without excessive loss of altitude or uncontrollable rolling or spinning tendencies.
(c) For limited elevator control stalls, it must be possible, when stalled from an excessive climb attitude, to recover without exceeding airspeed or acceleration limits.
Sec. 23.205 Stalls: critical engine inoperative.
(a) A multiengine airplane must have stall characteristics that prevent unintentional spin entry. This must be shown by performing the maneuver prescribed in paragraph (b) of this section, at the lowest practical altitude, with--
(1) The critical engine inoperative and its propeller in the normal inoperative position ;
(2) Landing gear extended, with the flaps (i) retracted and (ii) extended; and
(3) The remaining engines at full throttle or maximum continuous power.
(b) The maneuver required by paragraph (a) of this section is as follows: Establish a steady, curvilinear turn and, while maintaining a 15 degree bank (1), toward and (2) away from the inoperative engine, steadily increase the angle of attack with the elevator control until an uncontrollable downward pitching motion occurs. In performing this maneuver it must be possible to--
(1) Produce and correct roll by unreversed use of the lateral control until the airplane stalls; and
(2) Recover immediately to full flight control with wings level, from the stalled condition, by normal use of the controls, reducing power on the operating engines if desired without exceeding a 60 degree angle of bank.
Sec. 23.207 Stall warning.
There must be a clear and distinct stall warning with the flaps and landing gear in any position, both in straight and in turning flight. The stall warning must begin at a speed exceeding the stalling speed by not less than five, and not more than 10, miles per hour, and must continue until the stall occurs.
Spinning
Sec. 23.221 Spinning.
(a) Normal category. A single-engine, normal category airplane must be able to recover from a one-turn spin in not more than one additional turn, with the controls used in the manner normally used for recovery. In addition--
(1) For both the flaps-retracted and flaps-extended conditions, the applicable airspeed limit and positive limit maneuvering load factor may not be exceeded;
(2) There may be no excessive back pressure during the spin or recovery; and
(3) It must be impossible to obtain uncontrollable spins with any use of the controls.
For the flaps-extended condition, the flaps may be retracted during the recovery.
(b) Utility category. A utility category airplane must meet the requirements of paragraph (a) of this section or the requirements of paragraph (c) of this section.
(c) Acrobatic category. An acrobatic category airplane must be able to spin at least six turns, and must meet the following requirements:
(1) The airplane must recover from any point in a spin, not exceeding six turns with flaps retracted and one turn with flaps extended, in not more than one and one-half additional turns after normal recovery application of the controls.
(2) For both the flaps-retracted and flaps-extended conditions, the applicable airspeed limit and positive limit maneuvering load factor may not be exceeded. For the flaps-extended condition, the flaps may be retracted during recovery, if a placard is installed prohibiting intentional spins with flaps extended.
(3) It must be impossible to obtain uncontrollable spins with any use of the controls.
(d) Airplanes "characteristically incapable of spinning". If it is desired to designate an airplane as "characteristically incapable of spinning", this characteristic must be shown with--
(1) A weight five percent more than the highest weight for which approval is requested;
(2) A center of gravity at least three percent aft of the rearmost position for which approval is requested;
(3) An available elevator up-travel four degrees in excess of that to which the elevator travel is to be limited for approval; and
(4) An available rudder travel seven degrees, in both directions, in excess of that to which the rudder travel is to be limited for approval.
Ground and Water Handling Characteristics
Sec. 23.231 Longitudinal stability and control.
(a) A landplane may have no uncontrollable tendency to nose over in any reasonably expected operating condition, including rebound during landing or takeoff. Wheel brakes must operate smoothly and may not induce any undue tendency to nose over.
(b) A seaplane or amphibian may not have dangerous or uncontrollable porpoising characteristics at any normal operating speed on the water.
Sec. 23.233 Directional stability and control.
(a) There may be no uncontrollable ground or water looping tendency in 90 degree cross winds, up to a wind velocity of 0.2 VS0, at any speed at which the airplane may be expected to be operated on the ground or water.
(b) A landplane must be satisfactorily controllable, without exceptional piloting skill or alertness, in power-off landings at normal landing speed, without using brakes or engine power to maintain a straight path.
(c) The airplane must have adequate directional control during taxiing.
Sec. 23.235 Taxiing condition.
The shock-absorbing mechanism may not damage the structure of the airplane when the airplane is taxied on the roughest ground that may reasonably be expected in normal operation.
Sec. 23.239 Spray characteristics.
Spray may not dangerously obscure the vision of the pilots or damage the propellers or other parts of a seaplane or amphibian at any time during taxiing, takeoff, and landing.
Miscellaneous Flight Requirements
Sec. 23.251 Vibration and buffeting.
Each part of the airplane must be free from excessive vibration under any appropriate speed and power conditions up to at least the minimum value of VD allowed in Sec. 23.335. In addition, there may be no buffeting, in any normal flight condition, severe enough to interfere with the satisfactory control of the airplane, cause excessive fatigue to the crew, or result in structural damage. Stall warning buffeting within these limits is allowable.
Subpart C--Structure
General
Sec. 23.301 Loads.
(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
(b) Unless otherwise provided, the air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the airplane. These loads must be distributed to conservatively approximate or closely represent actual conditions.
(c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account.
(d) Simplified structural design criteria may be used if they result in design loads not less than those prescribed in Secs. 23.331 through 23.521. For conventional, single-engine airplanes with design weights of 6,000 pounds or less, the design criteria of Appendix A of this part are an approved equivalent of Secs. 23.331 through 23.399. If Appendix A is used, the entire Appendix must be substituted for the corresponding sections of this part.
Sec. 23.303 Factor of safety.
Unless otherwise provided, a factor of safety of 1.5 must be used.
Sec. 23.305 Strength and deformation.
(a) The structure must be able to support limit loads without detrimental, permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation.
(b) The structure must be able to support ultimate loads without failure for at least three seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the three second limit does not apply.
Sec. 23.307 Proof of structure.
(a) Compliance with the strength and deformation requirements of Sec. 23.305 must be shown for each critical load condition. Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable. In other cases, substantiating load tests must be made. Dynamic tests, including structural flight tests, are acceptable if the design load conditions have been simulated.
(b) Certain parts of the structure must be tested as specified in Subpart D of this part.
Flight Loads
Sec. 23.321 General.
(a) Flight load factors represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the airplane) to the weight of the airplane. A positive flight load factor is one in which the aerodynamic force acts upward, with respect to the airplane.
(b) Compliance with the flight load requirements of this subpart must be shown--
(1) At each critical altitude within the range in which the airplane may be expected to operate;
(2) At each weight from the design minimum weight to the design maximum weight; and
(3) For each required altitude and weight, for any practicable distribution of disposable load within the operating limitations specified in Secs. 23.1583 through 23.1589.
Sec. 23.331 Symmetrical flight conditions.
(a) The appropriate balancing horizontal tail load must be accounted for in a rational or conservative manner when determining the wing loads and linear inertia loads corresponding to any of the symmetrical flight conditions specified in Secs. 23.331 through 23.341.
(b) The incremental horizontal tail loads due to maneuvering and gusts must be reacted by the angular inertia of the airplane in a rational or conservative manner.
Sec. 23.333 Flight envelope.
(a) General. Compliance with the strength requirements of this subpart must be shown at any combination of airspeed and load factor on and within the boundaries of a flight envelope (similar to the one in paragraph (d) of this section) that represents the envelope of the flight loading conditions specified by the maneuvering and gust criteria of paragraphs (b) and (c) of this section respectively.
(b) Maneuvering envelope. Except where limited by maximum (static) lift coefficients, the airplane is assumed to be subjected to symmetrical maneuvers resulting in the following limit load factors:
(1) The positive maneuvering load factor specified in Sec. 23.337 at speeds up to VD;
(2) The negative maneuvering load factor specified in Sec. 23.337 at VC; and
(3) Factors varying linearly with speed from the specified value at VC to 0.0 at VD for the normal category, and -1.0 at VD for the acrobatic and utility categories.
(c) Gust envelope. Limit gust loads are the loads that would result when the airplane encounters the following symmetrical vertical gusts (assuming that gust load factors vary linearly between VC and VD):
(1) Positive (up) and negative (down) gusts of 30 feet per second nominal intensity at speeds up to VC.
(2) Positive and negative gusts of 15 feet per second at VD.
(d) Flight envelope.
Note: Point G need not be investigated when the supplementary condition specified in Sec. 23.369 is investigated.
Sec. 23.335 Design airspeeds.
The selected design airspeeds are equivalent airspeeds (EAS).
(a) Design cruising speed, VC. For VC, the following apply:
(1) VC (in miles per hour) may not be less than--
(i) 38  (for normal and utility category airplanes); and
(ii) 42  (for acrobatic category airplanes).
(2) For values of W/S more than 20, the numerical multiplying factors must be decreased linearly with W/S to a value of 33 where W/S =100. The required minimum value need not be more than 0.9 VH obtained at sea level.
(b) Design dive speed, VD. For VD, the following apply:
(1) With VC min the required minimum design cruising speed, VD (in miles per hour) may not be less than--
(i) 1.40 VC min (for normal category airplanes);
(ii) 1.50 VC min (for utility category airplanes); and
(iii) 1.55 VC min (for acrobatic category airplanes).
(2) For values of W/S more than 20, the numerical multiplying factors must be decreased linearly with W/S to a value of 1.35 at W/S = 100.
(c) Design maneuvering speed VA. For VA, the following applies:
(1) VA (in miles per hour) may not be less than VS  where--
(i) VS is a computed stalling speed with flaps retracted at the design weight, normally based on the maximum airplane normal force coefficients, CNA; and
(ii) n is the limit maneuvering load factor used in design.
(2) The value of VA need not exceed the value of VC used in design.
Sec. 23.337 Limit maneuvering load factors.
(a) The positive limit maneuvering load factor n may not be less than--
(1) 2.1+ for normal category airplanes, except that n need not be more than 3.8 nor may it be less than 2.5;
(2) 4.4 for utility category airplanes; or
(3) 6.0 for acrobatic category airplanes.
(b) The negative limit maneuvering load factor may not be less than--
(1) 0.4 times the positive load factor for the normal and utility categories; or
(2) 0.5 times the positive load factor for the acrobatic category.
(c) Maneuvering load factors lower than those specified in this section may be used if the airplane has design features that make it impossible to exceed these values in flight.
Sec. 23.341 Gust load factors.
In applying gust load requirements--
(a) The slope of the lift curve may be assumed to be that of the wing alone; and
(b) The gust load factors must be computed as follows:
n = 1 +
Where--
K = 1/2(W/S)1/4(for W/S<16 p.s.f.);
K = 1.33 -  (for W/S>16 p.s.f.);
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.
Sec. 23.345 High lift devices.
(a) If flaps or similar high lift devices to be used for takeoff, approach or landing are installed, the airplane, with the flaps fully deflected at VF, is assumed to be subjected to symmetrical maneuvers and gusts resulting in limit load factors within the range determined by--
(1) Maneuvering, to a positive limit load factor of 2.0; and
(2) Positive and negative gusts of 15 feet per second acting normal to the flight path in level flight.
(b) VF must be assumed to be not less than 1.4 VS or 1.8 VSF, whichever is greater, where--
VS is the computed stalling speed with flaps retracted at the design weight; and
VSF is the computed stalling speed with flaps fully extended at the design weight.
However, if an automatic flap load limiting device is used, the airplane may be designed for the critical combinations of airspeed and flap position allowed by that device.
(c) In designing the flaps and supporting structures, slipstream effects must be accounted for, as specified in paragraph (b) of Sec. 23.457.
(d) In determining external loads on the airplane as a whole, thrust, slipstream, and pitching acceleration may be assumed to be zero.
(e) The requirements of Secs. 23.175(c), 23.457, and this section, may be complied with separately or in combination.
Sec. 23.347 Unsymmetrical flight conditions.
The airplane is assumed to be subjected to the unsymmetrical flight conditions of Secs. 23.349 and 23.351. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner, considering the principal masses furnishing the reacting inertia forces.
Sec. 23.349 Rolling conditions.
The wing and wing bracing must be designed for the following loading conditions:
(a) Unsymmetrical wing loads appropriate to the category. Unless the following values result in unrealistic loads, the rolling accelerations may be obtained by modifying the symmetrical flight conditions in Sec. 23.333(d) as follows:
(1) For the acrobatic category, in conditions A and F, assume that 100 percent of the wing airload acts on one side of the plane of symmetry and 60 percent of this load acts on the other side.
(2) For the normal and utility categories, in condition A, assume that 100 percent of the wing airload acts on one side of the airplane and 70 percent of this load acts on the other side. For airplanes of more than 1,000 pounds design weight, the latter percentage may be increased linearly with weight up to 75 percent at 12,500 pounds.
(b) The loads resulting from the aileron deflections and speeds specified in Sec. 23.455, in combination with an airplane load factor of at least two thirds of the positive maneuvering load factor used for design. Unless the following values result in unrealistic loads, the effect of aileron displacement on wing torsion may be accounted for by adding the following increment to the basic airfoil moment coefficient over the aileron portion of the span in the critical condition determined in Sec. 23.333(d):
 cm=-0.01
where--
 cm is the moment coefficient increment; and
 is the down aileron deflection in degrees in the critical condition.
Sec. 23.351 Yawing conditions.
The airplane must be designed for yawing loads on the vertical tail surfaces resulting from the loads specified in Secs. 23.441 through 23.445.
Sec. 23.361 Engine torque.
(a) Each engine mount and its supporting structure must be designed for the effects of--
(1) The limit torque corresponding to takeoff power and propeller speed acting simultaneously with 75 percent of the limit loads from flight condition A; and
(2) The limit torque corresponding to maximum continuous power and propeller speed acting simultaneously with the limit loads from flight condition A.
(b) The limit torque is obtained by multiplying the mean torque by a factor of--
(1) 1.33 for engines with five or more cylinders; or
(2) Two, three, or four, for engines with four, three, or two cylinders, respectively.
(c) Engine torque effects need not be investigated for any other conditions.
Sec. 23.363 Side load on engine mount.
(a) Each engine mount and its supporting structure must be designed for a limit load factor in a lateral direction, for the side load on the engine mount, of not less than--
(1) 1.33, or
(2) One-third of the limit load factor for flight condition A.
(b) The side load prescribed in paragraph (a) of this section may be assumed to be independent of other flight conditions.
Sec. 23.365 Pressurized cabin loads.
For each pressurized compartment, the following apply:
(a) The airplane structure must be strong enough to withstand the flight loads combined with pressure differential loads from zero up to the maximum relief valve setting.
(b) The external pressure distribution in flight, and any stress concentrations, must be accounted for.
(c) If landings may be made, with the cabin pressurized, landing loads must be combined with pressure differential loads from zero up to the maximum allowed during landing.
(d) The airplane structure must be strong enough to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1.33, omitting other loads.
(e) If a pressurized cabin has two or more compartments separated by bulkheads or a floor, the primary structure must be designed for the effects of sudden release of pressure in any compartment with external doors or windows. This condition must be investigated for the effects of failure of the largest opening in the compartment. The effects of intercompartmental venting may be considered.
Sec. 23.369 Special conditions for rear lift truss.
(a) If a rear lift truss is used, it must be designed for conditions of reversed airflow at a design speed of--
V = 10  + 10 (m.p.h.).
(b) Either aerodynamic data for the particular wing section used, or a value of CL equaling -0.8 with a chordwise distribution that is triangular between a peak at the trailing edge and zero at the leading edge, must be used.
Control Surface and System Loads
Sec. 23.391 Control surface loads.
(a) The control surface loads specified in Secs. 23.397 through 23.459 are assumed to occur in the conditions described in Secs. 23.331 through 23.351.
(b) If allowed by the following sections, the values of control surface loading in Appendix B of this part may be used, instead of particular control surface data, to determine the detailed rational requirements of Secs. 23.397 through 23.459, unless these values result in unrealistic loads.
Sec. 23.395 Control system.
(a) Each flight control system and its supporting structure must be designed for loads corresponding to at least 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in Secs. 23.391 through 23.459. In addition, the following apply:
(1) The system limit loads need not exceed the higher of the loads that can be produced by the pilot and automatic devices operating the controls. However, autopilot forces need not be added to pilot forces. The system must be designed for the maximum effort of the pilot or autopilot, whichever is higher. In addition, if the pilot and the autopilot act in opposition, the part of the system between them may be designed for the maximum effort of the one that imposes the lesser load.
(2) The design must, in any case, provide a rugged system for service use, considering jamming, ground gusts, taxiing downwind, control inertia, and friction.
(b) A 125 percent factor on computed hinge moments must be used to design elevator, aileron, and rudder systems. However, a factor as low as 1.0 may be used if hinge moments are based on accurate flight test data, the exact reduction depending upon the accuracy and reliability of the data.
(c) Acceptable maximum and minimum pilot forces for elevator, aileron, and rudder controls are shown in Sec. 23.397(b). These pilot forces are assumed to act at the appropriate control grips or pads as they would in flight, and to react at the attachments of the control system to the control surface horns.
Sec. 23.397 Control system loads.
(a) General. In the control surface flight loading condition, the airloads on movable surfaces and the corresponding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph (b) of this section. In applying this criterion, the effects of control system boost and servo-mechanisms, and the effects of tabs must be considered. The automatic pilot effort must be used for design if it alone can produce higher control surface loads than the human pilot.
(b) The limit pilot forces. The limit pilot forces are as follows:
Control | Maximum forces for design weight W equal to or less than 5,000 pounds1 | Minimum forces2 |
| Aileron: | | |
| Stick----------------------------- | 67 pounds------------------------ | 40 pounds. |
| Wheel3-------------------------- | 53 D in-pounds4-------------- | 40 D in-pounds4. |
| Elevator: | | |
| Stick------------------------------ | 167 pounds---------------------- | 100 pounds. |
| Wheel---------------------------- | 200 pounds---------------------- | 100 pounds. |
| Rudder-------------------------- | 200 pounds---------------------- | 130 pounds. |
1 For design weight (W) more than 5,000 pounds, the specified maximum values must be increased linearly with weight to 1.18 times the specified values at a design weight of 12,500 pounds.
2 If the design of any individual set of control systems or surfaces makes these specified minimum forces inapplicable, values corresponding to the pertinent hinge moments obtained under Sec. 23.415, but not less than 0.6 of the specified minimum forces, may be used.
3 The critical parts of the aileron control system must also be designed for a single tangential force with a limit value of 1.25 times the couple force determined from the above criteria.
4 D=wheel diameter.
Sec. 23.399 Dual control system.
Each dual control system must be designed for the pilots operating in opposition, using individual pilot forces not less than--
(a) 0.75 times those obtained under Sec. 23.395; or
(b) The minimum forces specified in Sec. 23.397(b).
Sec. 23.405 Secondary control system.
Secondary controls, such as wheel brakes, spoilers, and tab controls, must be designed for the maximum forces that a pilot is likely to apply to those controls.
Sec. 23.407 Trim tab effects.
The effects of trim tabs on the control surface design conditions must be accounted for only where the surface loads are limited by maximum pilot effort. In these cases, the tabs are considered to be deflected in the direction that would assist the pilot. These deflections must correspond to the maximum degree of "out of trim" expected at the speed for the condition under consideration.
Sec. 23.409 Tabs.
Control surface tabs must be designed for the most severe combination of airspeed and tab deflection likely to be obtained within the flight envelope for any usable loading condition.
Sec. 23.415 Ground gust conditions.
(a) The control system must be investigated as follows for control surface loads due to ground gusts and taxiing downwind:
(1) If an investigation of the control system for ground gust loads is not required by subparagraph (2) of this paragraph, but the applicant elects to design a part of the control system for these loads, these loads need only be carried from control surface horns through the nearest stops or gust locks and their supporting structures.
(2) If pilot forces less than the minimums specified in Sec. 23.397(b) are used for design, the effects of surface loads due to ground gusts and taxiing downwind must be investigated for the entire control system according to the formula:
H = K c S q
where--
H = limit hinge moment (ft.-lbs.);
c = mean chord of the control surface aft of the hinge line (ft.);
S = area of control surface aft of the hinge line (sq. ft.);
q = dynamic pressure (p.s.f.) based on a design speed not less than 10
+10 (m.p.h.) except that the design speed need not exceed 60 m.p.h.; and
K = limit hinge moment factor for ground gusts derived in paragraph (b) of this section. (For ailerons and elevators, a positive value of K indicates a moment tending to depress the surface and a negative value of K indicates a moment tending to raise the surface).
(b) The limit hinge moment factor K for ground gusts must be derived as follows:
Surface | K | Position of controls |
| (a) Aileron---------------------------------- | 0.75 | Control column locked or lashed in mid-position. |
| (b) Aileron---------------------------------- | ±0.50 | Ailerons at full throw;
+ moment on one aileron,
- moment on the other. |
| (c) Elevator-------------------------------- | ±0.75 | (c) Elevator full up (-). |
| (d) Elevator-------------------------------- | -------- | (d) Elevator full down (+). |
| (e) Rudder--------------------------------- | ±0.75 | (e) Rudder in neutral. |
| (f) Rudder---------------------------------- | -------- | (f) Rudder at full throw. |
Horizontal Tail Surfaces
Sec. 23.421 Balancing loads.
(a) A horizontal tail balancing load is a load necessary to maintain equilibrium in any specified flight condition with no pitching acceleration.
(b) Horizontal tail surfaces must be designed for the balancing loads occurring at any point on the limit maneuvering envelope and in the flap conditions specified in Sec. 23.345. The distribution in figure 6 of Appendix B may be used.
Sec. 23.423 Maneuvering loads.
Each horizontal tail surface must be designed for maneuvering loads imposed by the following conditions:
(a) A sudden deflection of the elevator control, at VA, to (1) the maximum upward deflection, and (2) the maximum downward deflection, as limited by the control stops, or pilot effort, whichever is critical. The average loading of B23.11 of Appendix B and the distribution in figure 7 of Appendix B may be used.
(b) A sudden upward deflection of the elevator, at speeds above VA, followed by a downward deflection of the elevator, resulting in the following combinations of normal and angular acceleration:
Condition | Normal
acceleration (n) | Angular acceleration
(radian/sec.2) |
| Down load------------------------ | 1.0---------------------------------- |  nm (nm-1.5) |
| Up load---------------------------- | nm--------------------------------- |  nm(nm-1.5) |
where--
(1) nm=positive limit maneuvering load factor used in the design of the airplane; and
(2) V=initial speed in miles per hour.
The conditions in this paragraph involve loads corresponding to the loads that may occur in a "checked maneuver" (a maneuver in which the pitching control is suddenly displaced in one direction and then suddenly moved in the opposite direction), the deflections and timing avoiding exceeding the limit maneuvering load factor. The total tail load for both down and up load conditions is the sum of the balancing tail loads at V and the specified value of the normal load factor n, plus the maneuvering load increment due to the specified value of the angular acceleration. The maneuvering load increment in figure 2 of Appendix B and the distributions in figure 7 (for down loads) and in figure 8 (for up loads) of Appendix B may be used.
Sec. 23.425 Gust loads.
(a) Each horizontal tail surface must be designed for loads resulting from--
(1) Positive and negative gusts of 30 feet per second nominal intensity at VC, corresponding to the flight condition specified in Sec. 23.333(c), with flaps retracted; and
(2) Positive and negative gusts of 15 feet per second nominal intensity at VF , corresponding to the flight condition specified in Sec. 23.345(a)(2), with flaps extended and at VD corresponding with the flight conditions specified in Sec. 23.333(c)(2) with flaps retracted.
(b) The average loadings in figures 3 and 4 of Appendix B and the distribution in figure 8 of Appendix B may be used instead of the requirements of subparagraph (a)(1).
(c) When determining the total load on the horizontal tail for the conditions specified in paragraph (a) of this section, the initial balancing tail loads for steady unaccelerated flight at the pertinent design speeds VF, VC, and VD must first be determined. The incremental tail load resulting from the gusts must be added to the initial balancing tail load to obtain the total tail load.
(d) The incremental tail load due to the gust may be computed by the formula
 t = 0.1KUVStat 
where--
 t= the limit gust load increment on the tail in pounds;
K = gust coefficient K derived from Sec. 23.341;
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; and
Rw = aspect ratio of the wing.
Sec. 23.427 Unsymmetrical loads.
The maximum horizontal tail surface loading (load per unit area), as determined under Secs. 23.421 through 23.425, must be applied to the horizontal surfaces on one side of the plane of symmetry and the following percentage of that loading must be applied to the opposite side:
% =100-10 (n-1), where n is the specified positive maneuvering load factor
This value may not be more than 80 percent.
Vertical Tail Surfaces
Sec. 23.441 Maneuvering loads.
(a) At speeds up to VA, the vertical tail surfaces must be designed to withstand-
(1) A sudden displacement of the rudder control (with the airplane in unaccelerated flight with zero yaw) to the maximum deflection allowed by the control stops or by pilot strength, whichever is critical;
(2) A yaw angle of 15 degrees with the rudder fully deflected (except as limited by pilot strength) in the direction tending to increase the slip; and
(3) A yaw angle of 15 degrees with the rudder control maintained in the neutral position (except as limited by pilot strength).
(b) The average loading of B23.11 and figure 1 of Appendix B and the distribution in figures 7, 6, and 8 of Appendix B may be used instead of the requirements of subparagraphs (a)(1), (a)(2), and (a)(3), respectively.
(c) The yaw angles specified in paragraph (a)(3) of this section may be reduced if the yaw angle chosen for a particular speed cannot be exceeded in--
(1) Steady slip conditions;
(2) Uncoordinated rolls from steep banks; or
(3) Sudden failure of the critical engine with delayed corrective action.
Sec. 23.443 Gust loads.
(a) Vertical tail surfaces must be designed to withstand, in unaccelerated flight at VC, a gust of 30 feet per second nominal intensity normal to the plane of symmetry.
(b) The gust loading for that part of a vertical tail surface with a well defined leading edge must be computed by the formula
Where--
 =average limit unit pressure in pounds per square foot;
K=  , except that K may not be less than 1.0.
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; and
SV=vertical surface area in square feet.
A value of K obtained by rational determination may be used.
(c) The average loading in figure 5 and the distribution in figure 8 of Appendix B may be used.
Sec. 23.445 Outboard fins.
(a) If outboard fins are on the horizontal tail surface, the tail surfaces must be designed for the maximum horizontal surface load in combination with the corresponding loads induced on the vertical surfaces by endplate effects. These induced effects need not be combined with other vertical surface loads.
(b) If outboard fins extend above and below the horizontal surface, the critical vertical surface loading (the load per unit area as determined under Secs. 23.441 and 23.443) must be applied to--
(1) The part of the vertical surfaces above the horizontal surface with 80 percent of that loading applied to the part below the horizontal surface; and
(2) The part of the vertical surfaces below the horizontal surface with 80 percent of that loading applied to the part above the horizontal surface.
Ailerons, Wing Flaps, and Special Devices
Sec. 23.455 Ailerons.
(a) The ailerons must be designed for the loads to which they are subjected--
(1) In the neutral position during symmetrical flight conditions; and
(2) By the following deflections (except as limited by pilot effort), during unsymmetrical flight conditions:
(i) Sudden maximum displacement of the aileron control at VA. Suitable allowance may be made for control system deflections.
(ii) Sufficient deflection at VC, where VC is more than VA, to produce a rate of roll not less than obtained in subparagraph (2)(i).
(iii) Sufficient deflection at VD to produce a rate of roll not less than one-third of that obtained in subparagraph (2)(i).
(b) The average loading in B23.11 and figure 1 of Appendix B and the distribution in figure 9 of Appendix B may be used.
Sec. 23.457 Wing flaps.
(a) The wing flaps, their operating mechanisms, and their supporting structures must be designed for critical loads occurring in the flaps-extended flight conditions with the flaps in any position. However, if an automatic flap load limiting device is used, these components may be designed for the critical combinations of airspeed and flap position allowed by that device.
(b) The effects of propeller slipstream, corresponding to takeoff power, must be taken into account at not less than 1.4 VS, where VS is the computed stalling speed with flaps fully retracted at the design weight. For the investigation of slipstream effects, the load factor may be assumed to be 1.0.
Sec. 23.459 Special devices.
The loading for special devices using aerodynamic surfaces (such as slots and spoilers) must be determined from test data.
Ground Loads
Sec. 23.471 General.
The limit ground loads specified in this subpart are considered to be external loads and inertia forces that act upon an airplane structure. In each specified ground load condition, the external reactions must be placed in equilibrium with the linear and angular inertia forces in a rational or conservative manner.
Sec. 23.473 Ground load conditions and assumptions.
(a) The design landing weight (the maximum weight for landing conditions at the maximum descent velocity) may be used for structural design purposes only. Except as provided in paragraphs (b) and (c) of this section, this weight may not be less than the maximum weight.
(b) The design landing weight may be as low as 95 percent of the maximum weight if--
(1) The structural limit load values at the maximum weight are not exceeded at speeds up to takeoff speed over terrain as rough as that expected in service;
(2) The minimum fuel capacity is enough for at least one-half hour of operation at maximum continuous power plus the capacity equal to a fuel weight equal to the difference between the maximum weight and the design landing weight; and
(3) The operating limitations limit the takeoff weight to ensure that landing weights in normal operation do not exceed the design landing weight.
(c) The design landing weight of a multiengine airplane may be less than 95 percent of the maximum weight if--
(1) The airplane meets the one-engine-inoperative climb requirements of Sec. 23.67; and
(2) Instead of the corresponding requirements of this part, compliance is shown with the following requirements of Part 25 [New]:
(i) The ground load requirements of Secs. 25.471 and 25.473.
(ii) The landing gear requirements of Secs. 25.721 through 25.733.
(iii) The fuel jettisoning system requirements of Sec. 25.1001.
(d) The selected limit vertical inertia load factor at the center of gravity of the airplane for the ground load conditions prescribed in this subpart may not be less than that which would be obtained when landing with a descent velocity (V), in feet per second, equal to 4.4  , except that this velocity need not be more than 10 feet per second and may not be less than seven feet per second.
(e) Wing lift not exceeding two-thirds of the weight of the airplane may be assumed to exist throughout the landing impact and to act through the center of gravity. The ground reaction load factor may be equal to the inertia load factor minus the ratio of the above assumed wing lift to the airplane weight.
(f) Energy absorption tests (to determine the limit load factor corresponding to the required limit descent velocities) must be made under Sec. 23.725.
(g) No inertia load factor used for design purposes may be less than 2.67, nor may the limit ground reaction load factor be less than 2.0, unless these lower values will not be exceeded in taxiing at speeds up to takeoff speed over terrain as rough as that expected in service.
Sec. 23.477 Landing gear arrangement.
Sections 23.479 through 23.483, or the conditions in Appendix C, apply to airplanes with conventional arrangements of main and nose gear, or main and tail gear.
Sec. 23.479 Level landing conditions.
(a) For a level landing, the airplane is assumed to be in the following attitudes:
(1) For airplanes with tail wheels, a normal level flight attitude.
(2) For airplanes with nose wheels, attitudes in which--
(i) The nose and main wheels contact the ground simultaneously; and
(ii) The main wheels contact the ground and the nose wheel is just clear of the ground.
The attitude used in subdivision (i) of this subparagraph may be used in the analysis required under subdivision (ii) of this subparagraph.
(b) When investigating landing conditions, the drag components simulating the forces required to accelerate the tires and wheels up to the landing speed must be properly combined with the corresponding instantaneous vertical ground reactions, assuming wing lift and a tire-sliding coefficient of friction of 0.8. However, the drag loads may not be less than 25 percent of the maximum vertical ground reactions (neglecting wing lift).
(c) In determining the wheel spin-up loads for landing conditions, the method set forth in Appendix D or the arbitrary drag components in Appendix C must be used. However, if Appendix D is used, the 25 percent value for the minimum drag component must be used.
Sec. 23.481 Tail down landing conditions.
(a) For a tail down landing, the airplane is assumed to be in the following attitudes:
(1) For airplanes with tail wheels, an attitude in which the main and tail wheels contact the ground simultaneously.
(2) For airplanes with nose wheels, a stalling attitude, or the maximum angle allowing ground clearance by each part of the airplane, whichever is less.
(b) For airplanes with either tail or nose wheels, ground reactions are assumed to be vertical, with the wheels up to speed before the maximum vertical load is attained.
Sec. 23.483 One-wheel landing conditions.
For the one-wheel landing condition, the airplane is assumed to be in the level attitude and to contact the ground on one side of the main landing gear. In this attitude, the ground reactions must be the same as those obtained on that side under Sec. 23.479.
Sec. 23.485 Side load conditions.
(a) For the side load condition, the airplane is assumed to be in a level attitude with only the main wheels contacting the ground and with the shock absorbers and tires in their static positions.
(b) The limit vertical load factor must be 1.33, with the vertical ground reaction divided equally between the main wheels.
(c) The limit side inertia factor must be 0.83, with the side ground reaction divided between the main wheels so that--
(1) 0.5 (W) is acting inboard on one side; and
(2) 0.33 (W) is acting outboard on the other side.
Sec. 23.493 Braked roll conditions.
Under braked roll conditions, with the shock absorbers and tires in their static positions, the following apply:
(a) The limit vertical load factor must be 1.33.
(b) The attitudes and ground contacts must be those described in Sec. 23.479 for level landings.
(c) A drag reaction equal to the vertical reaction at the wheel multiplied by a coefficient of friction of 0.8 must be applied at the ground contact point of each wheel with brakes, except that the drag reaction need not exceed the maximum value based on limiting brake torque.
Sec. 23.497 Supplementary conditions for tail wheels.
In determining the ground loads on the tail wheel and affected supporting structures, the following apply:
(a) For the obstruction load, the limit ground reaction obtained in the tail down landing condition is assumed to act up and aft through the axle at 45 degrees. The shock absorber and tire may be assumed to be in their static positions.
(b) For the side load, a limit vertical ground reaction equal to the static load on the tail wheel, in combination with a side component of equal magnitude, is assumed. In addition--
(1) If a swivel is used, the tail wheel is assumed to be swiveled 90 degrees to the airplane longitudinal axis with the resultant ground load passing through the axle;
(2) If a lock, steering device, or shimmy damper is used, the tail wheel is also assumed to be in the trailing position with the side load acting at the ground contact point; and
(3) The shock absorber and tire are assumed to be in their static positions.
Sec. 23.499 Supplementary conditions for nose wheels.
In determining the ground loads on nose wheels and affected supporting structures, and assuming that the shock absorbers and tires are in their static positions, the following conditions must be met:
(a) For aft loads, the limit force components at the axle must be--
(1) A vertical component of 2.25 times the static load on the wheel; and
(2) A drag component of 0.8 times the vertical load.
(b) For forward loads, the limit force components at the axle must be--
(1) A vertical component of 2.25 times the static load on the wheel; and
(2) A forward component of 0.4 times the vertical load.
(c) For side loads, the limit force components at ground contact must be--
(1) A vertical component of 2.25 times the static load on the wheel; and
(2) A side component of 0.7 times the vertical load.
Sec. 23.505 Supplementary conditions for skiplanes.
In determining ground loads on skiplanes, and assuming that the airplane is resting on the ground with one main ski frozen at rest and the other main ski and the tail ski free to slide, a limit side force equal to P/3 must be applied at the most convenient point near the tail assembly, with--
(a) P being the static ground reaction on the tail ski; and
(b) A factor of safety of 1.0.
Water Loads
Sec. 23.521 Water load conditions.
(a) The structure of seaplanes and amphibians must be designed for water loads developed during takeoff and landing with the seaplane in any attitude likely to occur in normal operation at appropriate forward and sinking velocities under the most severe sea conditions likely to be encountered.
(b) Unless the applicant makes a rational analysis of the water loads, or uses the standards in ANC-3, Secs. 25.523 through 25.537 of this chapter apply.
(c) Floats certificated under Part 4a of this chapter before November 9, 1945, may be installed on airplanes that are designed under this part.
Emergency Landing Conditions
Sec. 23.561 General.
(a) The airplane, although it may be damaged in emergency landing conditions, must be designed as prescribed in this section to protect each occupant under those conditions.
(b) The structure must be designed to give each occupant every reasonable chance of escaping serious injury in a minor crash landing when--
(1) Proper use is made of belts or harnesses provided for in the design; and
(2) The occupant experiences the ultimate inertia forces shown in the following table:
Ultimate Inertia Forces
| Normal and utility categories | Acrobatic category |
| Upward---------------------------- | 3.0g | 4.5g |
| Forward---------------------------- | 9.0g | 9.0g |
| Sideward-------------------------- | 1.5g | 1.5g |
(c) Each airplane with retractable landing gear must be designed to protect each occupant in a landing--
(1) With the wheels retracted;
(2) With moderate descent velocity; and
(3) Assuming--
(i) An upward ultimate inertia force of 3 g; and
(ii) A coefficient of friction of 0.5 at the ground.
(d) If a turnover is reasonably probable, the structure must be designed to protect the occupants in a complete turnover, assuming--
(1) An upward ultimate inertia force of 3 g; and
(2) A coefficient of friction of 0.5 at the ground.
(e) Except as provided in Sec. 23.787 the supporting structure must be designed to restrain, under loads up to those specified in paragraph (b)(2) of this section, each item of mass that could injure an occupant if it came loose in a minor crash landing.
Fatigue Evaluation
Sec. 23.571 Pressurized cabin.
The strength, detail design, and fabrication of the pressure cabin structure must be evaluated under either of the following:
(a) A fatigue strength investigation, in which the structure is shown by analysis, tests, or both to be able to withstand the repeated loads of variable magnitude expected in service.
(b) A fail safe strength investigation, in which it is shown by analysis, tests, or both that catastrophic failure of the structure is not probable after fatigue failure, or obvious partial failure, of a principal structural element, and that the remaining structures are able to withstand a static ultimate load factor of 75 p |