CHAPTER 3. DESIGN AND CONSTRUCTION
Section 1. GENERAL
132. SECTION 23.629 FLUTTER.
This subject is covered in AC 23.629-1A.
133.-137. RESERVED.
Section 2. CONTROL SYSTEMS
138. SECTION 23.671 GENERAL. (RESERVED).
138A. SECTION 23.672 STABILITY AUGMENTATION AND AUTOMATIC AND POWER OPERATED SYSTEMS (RESERVED).
139. SECTION 23.677 TRIM SYSTEMS.
a. Qualitative Evaluation. Trim should be qualitatively evaluated during all phases of the flight test program. Cockpit control trim devices should be evaluated for smoothness, sense of motion, and ease of operation, accessibility, and visibility of the trim tab indicators (both day and night).
Ease in establishing and maintaining a trim condition should be evaluated.
b. Electric Trim Background. Electrically-actuated, manually-controlled trim systems have been certificated in several ways, depending on systems design. The simpler systems are tested for failure in flight. More sophisticated systems, which generally incorporate dual-wire, splitactuating switches, may require a dual failure to produce a runaway. Analysis of these systems discloses that one switch could fail closed and remain undetected until a failure occurred in the other switch or circuit to produce a runaway. This is still considered acceptable if the applicant provided a preflight test procedure that will detect such a dormant failure. Service experience dictates that evaluation of fail-safe trim systems by analysis alone is not acceptable and flight testing is required.
c. Explanation.
(1) Fault Analysis. A fault analysis should be evaluated for each trim system.
(2) Single Failure and Back-up System. For a system in which the fault analysis indicates a single failure will cause runaway, flight tests should be conducted. For a system with backup features, or a redundant system where multiple failures would be required for runaway, the certification team should determine the extent of the flight tests necessary after consideration of the fault analysis and determination of the probability and effect of runaway. In all cases, flight test evaluations should be conducted to determine functional system/airplane compatibility in accordance with § 23.1301.
(3) Failure. For the purpose of a fault analysis, a failure is the first fault obviously detectable by the pilot and should follow probable combinations of undetectable failures assumed as latent failures existing at the occurrence of the detectable failure. When an initial failure also causes other failures, the initial failure and the subsequent failures are considered 101
to constitute a single failure for purposes of fault analysis. Only independent failures may be introduced into the fault analysis to show multiple failure integrity.
(4) Failure Warning. The first indication a pilot has of a trim runaway is a deviation from the intended flight path, abnormal control movements, or a warning from a reliable failure warning system. The following time delays after pilot recognition are considered appropriate:
(a) Takeoff, approach, landing - 1 second.
(b) Climb, cruise, descent - 3 seconds.
(5) Second Set of Controls. If a set of controls and instruments are provided for a second crew member, multi-function systems disconnect or quick-disconnect/interrupt switches, as appropriate, should be provided for both crew members regardless of minimum crew.
d. Definitions.
(1) Disconnect Switch. A switch, which has the primary purpose of stopping all movement of the electric trim system, that is located within immediate reach and readily accessible to the pilot. A circuit breaker is not considered to be a disconnect switch.
(2) Quick-Disconnect/Interrupt Switch. A switch or device, which is located on the control wheel on the side opposite the throttles, or on the stick control, that momentarily interrupts all movement of the electric trim system and can be operated without moving the hand from its normal position on the control. The primary purpose of the switch is to stop all movement of the electric trim system.
e. Procedures.
(1) Quick-Disconnect or Interrupt Switch. With a quick-disconnect or interrupt switch, disconnect may be initiated after the delay times given in paragraph 139c(4).
(2) Disconnect Switch. With a disconnect switch, the time delays given in paragraph 139c(4) should be applied prior to corrective action by use of primary controls. In addition to these time delays, an appropriate reaction time to disconnect the systems should be added. When there are other switches in the immediate area of the quick-disconnect, a time increment should be added to account for identifying the switch.
(3) Loads. The loads experienced as a result of the malfunction should normally not exceed an envelope of 0 to +2g. The positive limit may be increased if analysis has shown that neither the malfunction nor subsequent corrective action would result in a load beyond limit load. In this case, careful consideration should be given to the delay time applied, since it may be more difficult for the pilot to reach the disconnect switch.
(4) High Speed Malfunctions. When high speed malfunctions are introduced at VNE or VMO/MMO, whichever is appropriate, the speed excursion, using the primary controls and any speed reduction controls/devices, should not exceed the demonstrated upset speed established under § 23.253 for airplanes with a VMO/MMO speed limitation and a speed midway between VNE and VD or those airplanes certified with a VNE limitation.
(5) Speed Limitations. The use of a reduction of VNE/VMO/MMO in complying with paragraph e(4) of this section is not considered acceptable, unless these new speeds represent limitations for the overall operation of the airplane.
(6) Forces. The forces encountered in the tests should conform to the requirements of § 23.143 for temporary and prolonged application. Also, see paragraph 45 of this AC.
To allow expansion of the 0 to 2g envelope, the FAA suggests a test procedure that incorporates both control restrained and unrestrained malfunctions. The following test matrix considers the probability of trim runaways, high airframe limit loads, control stick/wheel configuration and absence of an autopilot system. Because rudder trim can be adjusted without the pilot directly in the control loop (i.e., feet on the floor), restrained runaways for rudder trim are not considered acceptable. (See Table 1 in this section.)
TABLE 1. TRIM SYSTEMS REQUIREMENTS
M aximum Maximum Force M aximum Axis Time Load(g) Attitude (restrained and Rate of Force (unrestrained) Change unrestrained) C hange (unrestrained) (restrained) __________________________________________________________________________________ Pitch recognition structural +/-45 degrees 60 pounds 20 pounds/sec +3 seconds limits NTE
3.5g __________________________________________________________________________________ Roll recognition structural +/-90 degrees 30 pounds 10 pounds/sec +3 seconds limits __________________________________________________________________________________ Yaw recognition structural +/-30 degrees 150 pounds N/A +3 seconds limits (unrestrained only)
Note 1: Restrained means the pilot is in the control loop (hands on) and unrestrained means the pilot is not in the control loop (hands off).
Note 2: Trim systems with a monitor/limiter will be tested at a magnitude just below that required for monitor/limiter trip.
Note 3: NTE is Not to Exceed.
140. SECTION 23.679 CONTROL SYSTEM LOCKS.
This subject is covered in AC 23-17.
140A SECTION 23.691 ARTIFICIAL STALL BARRIER SYSTEM. (RESERVED)
141. SECTION 23.697 WING FLAP CONTROLS. (RESERVED)
142. SECTION 23.699 WING FLAP POSITION INDICATOR. (RESERVED)
143. SECTION 23.701 FLAP INTERCONNECTION.
This subject is covered in AC 23-17.
144.-153. RESERVED.
Section 3. LANDING GEAR
154. SECTION 23.729 LANDING GEAR EXTENSION AND RETRACTION SYSTEM.
This subject is covered in AC 23-17.
155. SECTION 23.735 BRAKES. (RESERVED)
156.-160. RESERVED.
Section 4. PERSONNEL AND CARGO ACCOMMODATIONS
161. SECTION 23.771 PILOT COMPARTMENT. (RESERVED)
162. SECTION 23.773 PILOT COMPARTMENT VIEW.
a. Pilot Position and View. For all evaluations, the pilot(s) should be seated at the intended design eye level as determined by an installed guide, if established. If an intended design eye level is not provided, the normal seating position should be used. The field of view that should remain clear should include the area specified in § 23.775(e).
b. External View. The external vision should be evaluated in all lighting and environmental conditions (day and night) with the airplane in all attitudes normally encountered. Attention to windshield distortion or refraction should especially be given to the view toward the approach and runway lights and the runway markings. Since glare and reflection often differ with the sun's inclination, consideration should be given to evaluating the cockpit at midday and in early morning or late afternoon. If the windshield is heated, evaluations should be conducted with heat on and off. Distortion and refraction should be so low as to prevent any unsafe condition, unusual eye strain or fatigue. "Safe operation," as used in § 23.773(a)(1), includes the ability to conduct straight ahead and circling approaches under all approved operating conditions, including operations in high humidity and icing conditions (if appropriate).
c. Night Approval. If night approval is requested, all lighting, both internal and external, should be evaluated in likely combinations and under expected flight conditions. Instrument lighting should be evaluated at night under a variety of ambient conditions, including night IFR.
Windshield/side window reflections that distract from traffic avoidance, landing approach and landing are not acceptable. Landing lights, strobes, beacons, and recognition lights should be evaluated to ensure no adverse reflections or direct impingement into the cockpit.
d. Defog/Defrost/Deice. The adequacy of the defog/defrost/deice systems should be evaluated under the following conditions:
(1) Extended cold soak at maximum altitudes and minimum temperatures. The airplane should be exposed to a cold environment appropriate to minimum expected temperatures.
The airplane should be also evaluated after remaining outside on a cold night.
(2) The airplane should be exposed to cold temperatures (cold soaked) and then descended into a warmer, moister air mass to assess ability to maintain a clear field of view. To properly evaluate internal fogging, the test airplane should be flown at night at high altitude for at least two hours (or until the windshield temperature stabilizes). Then, using proposed AFM 104
procedures, the airplane should be rapidly descended to an approach and landing in a high humidity area (recommend dewpoint of least 70 °F (21 °C)). If manual clearing by the pilot(s) is required, it should be "easily" accomplished by an average pilot. The applicant should provide any special equipment required to accomplish the manual clearing. Repeated immediate clearing after manually wiping the windshield would not seem to fit the "easily cleared" requirements. The "easily cleared" aspects should also be evaluated considering that the fogged windshield could frost under certain conditions. If manual clearing is required, pilot workload should be carefully evaluated if IFR approval is sought.
(3) Evaluations should be conducted in moderate rain, day and night (if approval is sought), takeoffs, landings, and taxi.
e. Two Pilot Airplanes. It is recommended that two pilot airplanes have pilot visibility in accordance with Society of Automotive Engineers (SAE) Aerospace Standard AS 580B, "Pilot Visibility from the Flight Deck Design Objectives for Commercial Transport Aircraft."
f. Cockpit Camera. An evaluation and documentation of the cockpit using a binocular camera is highly desirable.
162.A SECTION 23.775 WINDSHIELDS AND WINDOWS.
a. For commuter category airplanes, assuming loss of vision through any one panel in front of the pilot(s), either side panels or co-pilot panels, or both, may be used to show that continued safe flight and landing is possible using these panels only while remaining seated at a pilot(s) station.
For aircraft to be certified for IFR, the applicant should show that a safe landing can be demonstrated with IFR certified minimum visibility conditions.
163. SECTION 23.777 COCKPIT CONTROLS. (RESERVED)
163A. SECTION 23.785 SEATS, BERTHS, LITTERS, SAFETY BELTS AND SHOULDER HARNESSES.
a. Explanation. This subpart requires an approved seat for each occupant.
b. Procedures. Confirm that, when approved production seats are in place, the seats can be easily adjusted and will remain in a locked position.
164. SECTION 23.803 EMERGENCY EVACUATION.
This subject is covered in AC 20-118A.
165. SECTION 23.807 EMERGENCY EXITS.
AC 23-17 addresses this subject.
166. SECTION 23.831 VENTILATION.
a. Explanation. This subpart requires the carbon monoxide concentration not to exceed one part in 20,000 parts of air, which is 0.005 of 1 percent or 50 parts-per-million. A sample matrix for CO concentration is given in figure 166-1.
b. Procedures.
Test for Carbon Monoxide--
(1) Airplane may be at any convenient weight and c.g. position.
(2) Using a “CO” indicator reading instrument, record the values for the following tests:
Climb M.C. Power or Full Throttle Speed VREF Mixture Full Rich Cruise 75% M.C. Power Mixture Glide Engine(s) throttled Speed VREF Flaps up
Windows and/or Vents Partly open Closed Partly open Closed Partly open Closed
a. Maximum Reading (Cockpit):
(1) Along Windows and /or Vents
(2) Along Floor
(3) Front of Instrument Panel
(4) Front of Pilots Face
b. Maximum Reading (cabin):
(1) Front
(2) Center
(3) Rear
AUXILIARY
POWER UNIT
Installed? No Yes
HEATERS
Installed? No Yes OTHERS
c. With Tester Directly in Front of Unit While Unit is Operating
TABLE 166-1 – Sample of CO-concentration Matrix
167.-175. RESERVED.
Section 5. PRESSURIZATION
176. SECTION 23.841 PRESSURIZED CABINS.
AC 23-17 addresses this subject.
177. SECTION 23.843 PRESSURIZATION TESTS. (RESERVED)
178.-188. RESERVED.
CHAPTER 4. POWERPLANT
Section 1. GENERAL
189. SECTION 23.901 INSTALLATION. (RESERVED)
190. SECTION 23.903 ENGINES.
a. Explanation.
(1) Automatic Propeller Feathering Systems. All parts of the feathering device that are integral with the propeller or attached to it in a manner that may affect propeller airworthiness should be considered. The determination of airworthiness should be made on the following basis:
(i) The automatic propeller feathering system should not adversely affect normal propeller operation and should function properly under all temperatures, altitudes, airspeeds, vibrations, accelerations, and other conditions to be expected in formal ground and flight operation.
(ii) The automatic device should be demonstrated to be free from malfunctioning, which may cause feathering under any conditions other than those under which it is intended to operate. For example, it should not cause feathering during the following:
(A) Momentary loss of power; and
(B) Approaches with reduced throttle settings.
(iii) The automatic propeller feathering system should be capable of operating in its intended manner whenever the throttle control is in the normal position to provide takeoff power. No special operations at the time of engine failure should be necessary on the part of the crew in order to make the automatic feathering system operative.
(iv) The automatic propeller feathering installation should be such that not more than one propeller will be feathered automatically even if more than one engine fails simultaneously.
(v) The automatic propeller feathering installation should be such that normal operation may be regained after the propeller has begun to feather automatically.
(vi) The automatic propeller feathering installation should incorporate a switch or equivalent means to make the system inoperative. (Also see §§ 23.67 and 23.1501.)
(vii) If performance credit is given for the automatic propeller feathering system, there should be means provided to satisfactorily preflight check the system.
(viii) Some turbopropeller airplanes are equipped with some type of engine ignition system intended for use during flight in heavy precipitation conditions and for takeoff/landing on wet or slush-covered runways. The engine ignition system may be either automatic or continuous. The purpose of this system is to prevent or minimize the possibility of an engine flameout due to water ingestion. Compatibility with auto-feather systems should be ensured.
(2) Negative Torque Sensing Systems. (RESERVED).
b. Procedures.
(1) Automatic and Manual Propeller Feathering System Operational Tests.
(i) Tests should be conducted to determine the time required for the propeller to change from windmilling (with the propeller controls set for takeoff) to the feathered position at the takeoff speed determined in § 23.51.
(ii) The propeller feathering system should be tested at one engine inoperative climb airspeed. The configuration should be the following:
(A) Critical engine inoperative;
(B) Wing flaps retracted;
(C) Landing gear retracted; and
(D) Cowl flaps closed.
If the feathered propeller has a residual rotation, this has to be considered for aircraft performance.
(iii) The propeller should be tested in the actual configuration for an emergency descent.
A sufficient speed range should be covered to assure that any propeller rotation is not hazardous. In addition, the propeller should not inadvertently unfeather during these tests.
(iv) In order to demonstrate that the feathering system operates satisfactorily, propeller feather should be demonstrated throughout both the airspeed and the altitude envelope since engine failure may occur at any time. Propeller unfeathering, manually or automatically, need only be demonstrated up to the maximum one-engine-inoperative service ceiling or maximum airstart altitude, whichever is higher. Satisfactory propeller unfeathering should also be demonstrated after a 30-minute cold soak.
(2) Continued Rotation of Turbine Engines.
(i) Means should be provided to completely stop the rotation of turbine engines if continued rotation would cause a hazard to the airplane. Devices such as feathering propellers, brakes, doors, or other means may be used to stop turbine engine rotation.
(ii) If engine induction air duct doors or other types of brakes are provided to control engine rotation, no single fault or failure of the system controlling engine rotation should cause the inadvertent travel of the doors toward the closed position or the inadvertent energizing of braking means unless compensating features are provided to ensure that engine failure or a critical operating condition will not occur. Such provisions should be of a high order of reliability, and the probability should be remote that doors or brakes will not function normally on demand.
(3) Engine Operation with Automatic Propeller Control System Installed.
(i) When an automatic control system for simultaneous r.p.m. control of all propellers is installed, it should be shown that no single failure or malfunction in this system or in the engine controlling this system will cause the following:
(A) The allowable engine overspeed for this condition to be exceeded at any time;
and
(B) A loss of thrust that will cause the airplane to fail to meet the requirements of §§ 23.51 through 23.77 if such system is certificated for use during takeoff and climb. This should be shown for all weights and altitudes for which certification is desired. A period of 5 seconds should be allowed from the time the malfunction occurs to the initial motion of the cockpit control for corrective action taken by the crew.
(ii) Compliance with this policy may be shown by analysis, flight demonstration, or a combination thereof.
c. Restart Envelope.
(1) Explanation. The applicant should propose a practicable airstart envelope wherein satisfactory inflight engine restarts may be accomplished as required by the code. Airstarts should be accomplished satisfactorily at critical combinations of airspeed and altitude.
During these tests, normally, time history data showing airspeed, altitude, r.p.m., exhaust temperature, and so forth, are obtained for inclusion in the Type Inspection Report.
The airstart envelope should be included in the limitations section of the AFM, and the procedures used to restart the engine(s) should be contained in the emergency or abnormal procedures section of the AFM.
Results of restart tests completed by the engine manufacturer on the same type of engine in an altitude test facility or flying test bed, if available. The experience accumulated in other aircraft with the same engine and engine installation, may be taken into account if justified.
(2) Procedures. To establish the required envelope of altitude and airspeed, sufficient flight tests should be made as follows:
(i) From sea-level to the maximum declared restarting altitude in all appropriate configurations likely to affect restarting, including the emergency descent configuration;
(ii) From the minimum to the maximum declared airspeed at all altitudes up to the maximum declared engine restarting altitude. The airspeed range of the declared restart envelope normally should cover at least 30 kts, but should be adapted to the type of airplane.
The tests should include the effect on engine restarting performance of delay periods between engine shutdown and restarting as follows:
(iii) Up to two minutes; and
(iv) At least until the engine oil temperature is stabilized at its approximate cold soak value.
191. SECTION 23.905 PROPELLERS.
Included in § 23.903 material. See paragraph 190 of this AC.
192. SECTION 23.909 TURBO SUPERCHARGERS.
AC 23-16 addresses this subject.
192.A. SECTION 23.925 PROPELLER CLEARANCE. (RESERVED)
193. SECTION 23.929 ENGINE INSTALLATION ICE PROTECTION.
a. Explanation. This regulation requires that propellers and other components of the complete engine installation, such as oil cooling inlets, generator cooling inlets, and so forth, function satisfactorily and operate properly without an appreciable and unacceptable loss of power when the applicant requests approval for flight in icing conditions. An unacceptable loss of power may depend on the kind of aircraft and the power available. For details see AC 23.1419-2. See § 23.1093 for induction system ice protection requirements.
b. Procedures. Each propeller and other components of the complete installation that is to be approved for flight in icing conditions should be evaluated under the icing conditions specified in Part 25, Appendix C. If the propellers are equipped with fluid-type deicers, the flow test should be conducted starting with a full tank of fluid and operated at maximum flow for a time period found operationally suitable. The operation should be checked at all engine speeds and powers.
194. SECTION 23.933 REVERSING SYSTEMS.
a. Explanation. Self-explanatory.
b. Procedures. Reversing system installations may be approved provided the following is acceptable:
(1) Exceptional pilot skill should not be required in taxiing or any condition in which reverse thrust is to be used.
(2) Necessary operating procedures, operating limitations, and placards are established.
(3) The airplane control characteristics are satisfactory with regard to control forces encountered, and buffeting should not cause structural damage.
(4) The directional control is adequate using normal piloting skill.
(5) A determination is made that no dangerous condition is encountered in the event of sudden failure of one engine in any likely operating condition.
(6) The operating procedures and airplane configuration are such as to provide reasonable safeguards against serious structural damage to parts of the airplane due to the reverse airflow.
(7) It is determined that the pilot's vision is not dangerously obscured under normal operating conditions on dusty or wet runways and where light snow is on the runway.
(8) It is determined that the pilot's vision is not dangerously obscured by spray due to reverse airflow under normal water operating conditions with seaplanes.
(9) The procedure and mechanisms for reversing should provide a reverse idle setting such that without requiring exceptional piloting skill at least the following conditions are met:
(i) Sufficient power is maintained to keep the engine running at an adequate speed to prevent engine stalling during and after the propeller reversing operation.
(ii) The propeller does not overspeed during and after the propeller reversing operation.
(10) The engine cooling characteristics should be satisfactory in any likely operating condition.
(11) If using ground idle for disking drag credit on landing distance, the ground idle position of the power levers should be identified with a gate or a detent with satisfactory tactile feel (reference paragraph 27a(7) of this AC).
(12) If compliance with § 23.933(a)(1)(ii) is intended to be shown by flight tests, any possible position of any one thrust reverser has to be assumed.
195. SECTION 23.939 POWERPLANT OPERATING CHARACTERISTICS.
a. Explanation. Self-explanatory.
b. Procedures.
(1) Stall, Surge, Flameout Tests. For turbine engines, tests should be conducted to determine that stall, surge, and flameout will not occur, to a hazardous degree, on any engine during acceleration and deceleration throughout the normal flight envelope of the airplane. This would include tests throughout the approved altitude range and throughout the airspeed range from VS to VMO/MMO using sideslip angles appropriate to the individual airplane. For normal category multiengine airplanes, an appropriate sideslip angle is generally considered to be approximately one ball width on a standard slip-skid indicator. The low airspeed tests should be accomplished at light weight and with gear and flaps extended to further reduce the stall speed. Tests need not be accomplished with gear and flaps extended at airspeeds above which extension is prohibited in the AFM. At the conditions mentioned above, the effects of engine bleed air off and on and engine ice protection systems off and on should be investigated.
(2) Throttle Techniques. With the engine stabilized at maximum continuous power, rapidly retard the throttle to the flight idle position. Before the engine reaches idle power or r.p.m., rapidly advance the throttle to maximum continuous power. Repeat this process, except begin with the engine stabilized at flight idle power. Rapid throttle movement is generally defined as one which results in the throttle moving from maximum continuous power to flight idle, or vice versa, in not more than 0.5 seconds.
196. SECTION 23.943 NEGATIVE ACCELERATION.
a. Explanation. Tests should be conducted to show that no hazardous malfunction occurs under negative accelerations within the flight envelope. A hazardous malfunction in this case usually is considered to be one which causes a loss or sustained malfunction of the engine, or improper operation of the engine accessories or systems.