# AC 23-17A: Systems and Equipment Guide for Certification of Part 23 Airplanes - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Historical - Version in force: [AC 23-17C](https://openairdocs.org/ac/23-17), approved November 17, 2011. This is a historical version - Approved: June 27, 2002 - Last reviewed: August 13, 2002 - Family: AC 23-17 - Cancels: [AC 23-17](https://openairdocs.org/ac/23-17) - Cancelled by: [AC 23-17B](https://openairdocs.org/ac/23-17B) - 14 CFR parts: 23 - 14 CFR sections: [23.671](https://www.ecfr.gov/current/title-14/section-23.671) - Pages: 246 - PDF: https://openairdocs.org/files/sha256/86/df/86df1632bfb58ba940e9ff2be64f2cbb27da923133c12333a8dee6f71276cb06 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/B0BC92E69A4362C386256BF7006C6823.0001 - DRS record last changed: November 21, 2025 - Page: https://openairdocs.org/ac/23-17A ## Versions - [AC 23-17C](https://openairdocs.org/ac/23-17) (Current), approved Nov 17, 2011, 371 pages - [AC 23-17B](https://openairdocs.org/ac/23-17B) (Historical), approved Apr 12, 2005, 342 pages - [AC 23-17A](https://openairdocs.org/ac/23-17A) (Historical), approved Jun 27, 2002, 246 pages - [AC 23-17](https://openairdocs.org/ac/23-17?v=B6D66E325E7EE39286256BF2004AB326) (Historical), approved Apr 25, 2000, 197 pages ## Text Each page's text from the PDF; open the PDF at a page with #page=N. Some pages are OCR from the scanned FAA advisory circular. OCR can misread letters and numbers. Those pages are marked (OCR). ### Page 1 Subject: SYSTEMS AND EQUIPMENT GUIDE FOR CERTIFICATION OF PART 23 AIRPLANES Date: 6/27/02 Initiated By: ACE-100 AC No: 23-17A Change: 1. PURPOSE. This advisory circular (AC) sets forth an acceptable means, but not the only means, of showing compliance with Title 14 Code of Federal Regulations (14 CFR) part 23 for the certification of systems and equipment in normal, utility, acrobatic, and commuter category airplanes. This AC applies to Subpart D from § 23.671 and Subpart F. This AC both consolidates existing policy documents, and certain AC's that cover specific paragraphs of the regulations, into a single document and adds new guidance. Material in this AC is neither mandatory nor regulatory in nature and does not constitute a regulation. 2. CANCELLATION. The following AC is canceled: AC 23-17, Systems and Equipment Guide for Certification of Part 23 Airplanes. 3. BACKGROUND. In 1968, the Federal Aviation Administration (FAA) instituted an extensive review of the airworthiness standards of part 23. Since then, the regulations have been amended through Amendment 23-53. These amendments have changed most of the sections of part 23. This document is intended to provide guidance for the original issue of part 23 and the various amendments. This version of the advisory circular covers policy available through June 30, 2001. Policy that became available after June 30, 2001, will be covered in future amendments to the advisory circular. 4. APPLICABILITY. This AC is applicable only to the original applicant seeking issuance of a Type Certificate (TC), an Amended TC, or a Supplemental Type Certificate (STC) for the initial approval of the new type design or a change in the approved type design. This material is not to be construed as having any legal status and should be treated accordingly. This version of the advisory circular covers policy available through June 30, 2001. Policy that became available after that date will be covered in future amendments to the advisory circular. ### Page 2 ii 5. PARAGRAPHS KEYED TO PART 23. Each paragraph has the applicable part 23 amendment shown in the title. As part 23 changes occur, the appropriate revisions will be made to the affected paragraphs of this AC. 6. RELATED PUBLICATIONS. These documents are provided as a quick reference source of documents that are acceptable for use in 14 CFR part 23 certification programs/projects. a. Free Orders and AC's Copies of current publications of the following free Orders and AC’s listed below can be obtained from the U.S. Department of Transportation, Subsequent Distribution Office, Ardmore East Business Center, 3341 Q 75th Avenue, Landover, MD 20785: The website where these orders and advisory circulars can be found is www.faa.gov/. FAA Order 8110.4B, Type Certification. FAA Order 8100.5, Aircraft Certification Directorate Procedures. AC 20-30B, Aircraft Position Light and Anticollision Light Installation. AC 20-36S, Index of Articles Certified Under the Technical Standard Order System. AC 20-41A, Substitute Technical Standard Order (TSO) Aircraft Equipment. AC 20-42C, Hand Fire Extinguishers for Use in Aircraft. AC 20-67B, Airborne VHF Communications Equipment Installations. AC 20-73, Aircraft Ice Protection. AC 20-74, Aircraft Position and Anticollision Light Measurements. AC 20-112, Airworthiness and Operational Approval of Airborne Systems to be Used in Lieu of a Ground Proximity Warning System (GPWS). AC 20-115B, Radio Technical Commission for Aeronautics, Inc., Document RTCA/DO- 178B. AC 20-118A, Emergency Evacuation Demonstration. AC 20-121A, Airworthiness Approval of Airborne Loran-C Navigation Systems for Use in the U.S. National Airspace System (NAS). ### Page 3 iii AC 20-124, Water Ingestion Testing for Turbine Powered Airplanes. AC 20-TCAS (Draft), Airworthiness Approval and Operational Use of Traffic Alert and Collision Avoidance System (TCAS I). AC 20-128A, Design Considerations for Minimizing Hazards Caused by Uncontained Turbine Engine and Auxiliary Power Unit Rotor Failure. AC 20-131A, Airworthiness and Operational Approval of Traffic Alert and Collision Avoidance Systems (TCAS II) and Mode S Transponders. AC 20-136, Protection of Aircraft Electrical/Electronic Systems Against the Indirect Effects of Lightning. AC 20-138, Airworthiness Approval of Global Positioning System (GPS) Navigation Equipment for Use as a VFR and IFR Supplemental Navigation System. AC 21-16D, Radio Technical Commission for Aeronautics (RTCA) Document DO-160D. AC 21-25A, Approval of Modified Seats and Berths Initially Approved Under a Technical Standard Order. AC 21-34, Shoulder Harness-Safety Installations. AC 23-2, Flammability Tests. AC 23-18, Installation of Terrain Awareness and Warning System (TAWS) Approved for Part 23 Airplanes. AC 23.143-1, Ice Contaminated Tailplane Stall. AC 23.562-1, Dynamic Testing of Part 23 Airplane Seat/Restraint Systems and Occupant Protection. AC 23.1309-1C, Equipment, Systems, and Installations in Part 23 Airplanes. AC 23.1311-1A, Installation of Electronic Display Instrument Systems in Part 23 Airplanes. AC 23.1419-2A, Certification of Part 23 Airplanes for Flight in Icing Conditions. AC 23-28, Airframe Guide for Certification of Part 23 Airplanes. ### Page 4 iv AC 25-11, Transport Category Airplane Electronic Display Systems. AC 90-79, Recommended Practices and Procedures for the Use of Electronic Long-Range Navigation. AC 120-31A, Operational and Airworthiness Approval of Airborne Omega Radio Navigation Systems as a Means of Updating Self-Contained Navigation Systems. AC 120-37, Operational and Airworthiness Approval of Airborne Omega Radio Navigation Systems as a Sole Means of Long-Range Navigation Outside the United States. AC 121-13, and Change 1 and 2, Self-Contained Navigation Systems (Long Range). Copies of current publications of the following “for sale” AC’s may be purchased from the Superintendent of Documents, P. O. Box 371954, Pittsburgh, PA 15250-7954; make check or money order payable to the Superintendent of Documents: AC 20-88A, Guidelines on the Marking of Aircraft Powerplant Instruments (Displays). AC 20-101C, Airworthiness Approval of Omega/VLF Navigation Systems for Use in the United States NAS and Alaska. AC 21.303-2H, Announcement of Availability: Parts Manufacturer Approvals1992 (Microfiche). AC 23-16, Powerplant Guide for Certification of Part 23 Airplanes. AC 23-8A and Change 1, Flight Test Guide for Certification of Part 23 Airplanes. AC 43.13-1B, Acceptable Methods, Techniques, and PracticesAircraft Inspection and Repair. AC 43.13-2A and Change 2, Acceptable Methods, Techniques, and PracticesAircraft Alterations (includes Change 1). NOTE: Republishing these AC documents as a part of this AC was not considered to be the best utilization of FAA resources. ### Page 5 v b. Industry Documents (1) To obtain a copy of the Technical Standard Orders (TSO’s), write to the U.S. Department of Transportation, Subsequent Distribution Office, Ardmore East Business Center, 3341 Q 75th Avenue, Landover, MD 20785, or use the website www.faa.gov/: TSO-C9c, Automatic Pilots. TSO-C62d, Aircraft Tires. TSO-C22g, Safety Belts. TSO-C26c, Aircraft Wheels and Wheel-Brakes Assemblies, with Addendum I. TSO-C39b, Aircraft Seats and Berths. TSO-C55, Fuel and Oil Quantity Instruments (For Reciprocating Engine Aircraft). TSO-C114, Torso Restraint Systems. TSO-C151a, Terrain Awareness and Warning System. (2) The RTCA documents listed below are available from RTCA, Inc., Suite 1020, 1140 Connecticut Avenue, NW, Washington, DC 20036-4001: RTCA/DO-160D, Environmental Test Conditions and Test Procedures for Airborne Equipment. RTCA/DO-178B, Software Considerations in Airborne Systems and Equipment Certification. (3) SAE stands for Society of Automotive Engineers. The SAE documents listed below are available from the Society of Automotive Engineers, Inc., 400 Commonwealth Drive, Warrendale, PA 15096-0001: ARP 597C, Wheels and Brakes, Supplementary Criteria Design for EnduranceCivil Transport Aircraft. ARP 813A, Maintainability Recommendations for Aircraft Wheels and Brakes. ARP 1619A, Replacement and Modified Brakes and Wheels. AIR 1064B, Brake Dynamics. ### Page 6 vi AS 1145A, Aircraft Brake Temperature Monitor System. SAE J384, Motor Vehicle Seat Belt Anchorage’s Test Procedure. SAE Recommended Practice, 1979 SAE Handbook, Volume 2, pages 33.08-33.09. SAE ARP 5412, Aircraft Lightning Environment and Related Test Waveforms. SAE ARP 5413, Certification of Aircraft Electrical/Electronic Systems for the Indirect Effects of Lightning, (it is being converted to an AC 20-136A). SAE ARP 5414, Aircraft Lightning Zoning. SAE ARP 5475, Abuse Load Testing for In-Seat Deployable Video Systems. (4) The Underwriter’s Laboratories (UL), Inc., document listed below can be obtained from Global Engineering Documents, 15 Inverness Way East, Englewood, CO 80112: UL 1418, Implosion Protected Cathode Ray Tubes for Television Type Appliances, Revised 1992. c. Joint Aviation Authorities (JAA) Documents Joint Aviation Requirements (JAR's) and guidance materials, Advisory Circular, Joint (ACJ's), are available by subscription from the JAA website at www.jaa.nl/ then click on JAR's. d. NASA Document NSS 1740.15, NASA Safety Standard for Oxygen and Oxygen Systems, Revised January 1996. s/ Michael Gallagher Manager, Small Airplane Directorate Aircraft Certification Service ### Page 7 vii CONTENTS Subpart D—Design and Construction CONTROL SYSTEMS Section PAGE 23.671 General … 1 23.672 Stability augmentation and automatic and power-operated systems … 2 23.673 Primary flight controls … 3 23.675 Stops … 4 23.677 Trim systems … 5 23.679 Control system locks … 7 23.681 Limit load static tests … 8 23.683 Operation tests … 9 23.685 Control system details … 18 23.687 Spring devices … 19 23.689 Cable systems … 20 23.691 Artificial stall barrier system … 21 23.693 Joints … 23 23.697 Wing flap controls … 24 23.699 Wing flap position indicator … 25 23.701 Flap interconnection … 26 23.703 Takeoff warning system … 30 **LANDING GEAR** 23.721 General … 31 23.723 Shock absorption tests … 32 23.725 Limit drop tests … 33 23.726 Ground load dynamic tests … 34 23.727 Reserve energy absorption drop test … 35 23.729 Landing gear extension and retraction system … 36 23.731 Wheels … 40 23.733 Tires … 41 23.735 Brakes … 54 23.737 Skis … 71 23.745 Nose/tail wheel steering … 72 ### Page 8 viii FLOATS AND HULLS Section PAGE 23.751 Main float buoyancy … 73 23.753 Main float design … 74 23.755 Hulls … 75 23.757 Auxiliary floats … 76 **PERSONNEL AND CARGO ACCOMMODATIONS** 23.771 Pilot compartment … 77 23.773 Pilot compartment view … 78 23.775 Windshields and windows … 79 23.777 Cockpit controls … 81 23.779 Motion and effect of cockpit controls … 84 23.781 Cockpit control knob shape … 85 23.783 Doors … 86 23.785 Seats, berths, litters, safety belts, and shoulder harnesses … 89 23.787 Baggage and cargo compartments … 102 23.791 Passenger information signs … 105 23.803 Emergency evacuation … 106 23.805 Flight crew emergency exits … 107 23.807 Emergency exits … 108 23.811 Emergency exit marking … 116 23.812 Emergency lighting … 117 23.813 Emergency exit access … 118 23.815 Width of aisle … 119 23.831 Ventilation … 120 **PRESSURIZATION** 23.841 Pressurized cabins … 121 23.843 Pressurization tests … 125 **FIRE PROTECTION** 23.851 Fire extinguishers … 126 23.853 Passenger and crew compartment interiors … 127 23.855 Cargo and baggage compartment fire protection … 131 23.859 Combustion heater fire protection … 132 23.863 Flammable fluid fire protection … 133 23.865 Fire protection of flight controls, engine mounts, and other flight structure … 134 ### Page 9 ix ELECTRICAL BONDING AND LIGHTNING PROTECTION Section PAGE 23.867 Electrical bonding and protection against lightning and static electricity … 136 **MISCELLANEOUS** 23.871 Leveling means … 138 Subpart F—Equipment **GENERAL** 23.1301 Function and installation … 139 23.1303 Flight and navigation instruments … 145 23.1305 Powerplant instruments … 152 23.1307 Miscellaneous equipment … 161 23.1309 Equipment, systems, and installations … 162 INSTRUMENTS: INSTALLATION 23.1311 Electronic display instrument systems … 165 23.1321 Arrangement and visibility … 166 23.1322 Warning, caution, and advisory lights … 167 23.1323 Airspeed indicating system … 168 23.1325 Static pressure system … 169 23.1326 Pitot heat indication systems … 170 23.1327 Magnetic direction indicator … 171 23.1329 Automatic pilot system … 173 23.1331 Instruments using a power source … 185 23.1335 Flight director systems … 188 23.1337 Powerplant instruments installation … 189 **ELECTRICAL SYSTEMS AND EQUIPMENT** 23.1351 General … 190 23.1353 Storage battery design and installation … 191 23.1357 Circuit protective devices … 194 23.1359 Electrical system fire protection … 196 23.1361 Master switch arrangement … 197 23.1365 Electric cables and equipment … 198 23.1367 Switches … 199 ### Page 10 x LIGHTS Section PAGE 23.1381 Instrument lights … 200 23.1383 Taxi and landing lights … 201 23.1385 Position light system installation … 202 23.1387 Position light system dihedral angles … 203 23.1389 Position light distribution and intensities … 204 23.1391 Minimum intensities in the horizontal plane of position lights … 205 23.1393 Minimum intensities in any vertical plane of position lights … 206 23.1395 Maximum intensities in overlapping beams of position lights … 207 23.1397 Color specifications … 208 23.1399 Riding light … 209 23.1401 Anticollision light system … 210 **SAFETY EQUIPMENT** 23.1411 General … 211 23.1413 Safety Belts and Harnesses [Removed] … 212 23.1415 Ditching equipment … 213 23.1416 Pneumatic de-icer boot system … 214 23.1419 Ice protection … 215 **MISCELLANEOUS EQUIPMENT** 23.1431 Electronic equipment … 217 23.1435 Hydraulic systems … 219 23.1437 Accessories for multiengine airplanes … 220 23.1438 Pressurization and pneumatic systems … 221 23.1441 Oxygen equipment and supply … 222 23.1443 Minimum mass flow of supplemental oxygen … 224 23.1445 Oxygen distribution system … 225 23.1447 Equipment standards for oxygen dispensing units … 226 23.1449 Means for determining use of oxygen … 228 23.1450 Chemical oxygen generators … 229 23.1451 Fire protection for oxygen equipment … 230 23.1453 Protection of oxygen equipment from rupture … 231 23.1457 Cockpit voice recorders … 232 23.1459 Flight recorders … 233 23.1461 Equipment containing high energy rotors … 234 ### Page 11 SYSTEMS AND EQUIPMENT GUIDE FOR CERTIFICATION OF PART 23 AIRPLANES Subpart D—Design and Construction **CONTROL SYSTEMS** 23.671 General No FAA policy is available as of June 30, 2001. JAA ACJ 23.671 is acceptable for FAA certification. ### Page 12 23.672 Stability augmentation and automatic and power-operated systems Amendment 23-45 and Subsequent This rule is applicable only if the system is required to show compliance with the flight characteristic requirements of part 23. ### Page 13 23.673 Primary flight controls No policy available as of June 30, 2001. ### Page 14 23.675 Stops No policy available as of June 30, 2001. ### Page 15 23.677 Trim systems Original Issue and Subsequent The trim system should prevent inadvertent, improper or abrupt trim operation. The direction of trim movement and its relation to its range of adjustment should be designed to prevent confusion. Trim devices should be designed to continue normal operation with one failure of any connecting or transmitting element in the primary flight control system for (1) longitudinal trim in a single-engine airplane, and (2) longitudinal and directional trim in multiengine airplanes. Amendment 23-7 and Subsequent The amendment requires there be adequate control for safe flight and landing (rather than to “continue normal operation”) using the trim devices following the failure of a connecting/transmitting element in the primary controls. Thus, the control system element failure must not cause a failure of the trim system. Failures of the trim system must not prevent safe flight and landing. Amendment 23-34 and Subsequent Probable powered trim runaways should be demonstrated for all part 23 airplanes so equipped. See AC 23-8A, Change 1, Flight Test Guide for Certification of part 23 Airplanes, for the procedure. Even if trim runaways have been determined to be improbable using the guidance in AC 23.1309-1C, Equipment, Systems, and Installations in Part 23 Airplanes, appropriate trim runaway demonstrations in all axes are required to demonstrate that the airplane has no unsafe features. The FAA has accepted demonstration of control restrained trim runaways during malfunction testing for systems without a monitor/limiter regardless of the reliability and those with a monitor/limiter whose reliability is less than extremely improbable. However, the FAA has determined this procedure is not acceptable in itself for failure conditions shown to be less than extremely improbable. To allow expansion of the 0 to 2g envelope, as specified in AC 23-8A, 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), ### Page 16 6 restrained runaways for rudder trim are not considered acceptable. (See Table 1 in this section.) TABLE 1. TRIM SYSTEMS REQUIREMENTS Maximum Maximum Force Maximum Axis Time Load(g) Attitude (restrained and Rate of Force (unrestrained) Change unrestrained) Change (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. ### Page 17 23.679 Control system locks Original Issue and Subsequent Section 23.679(a) of part 23 and § 3.341(a) of the CAR require that if there is a device to lock the control system, there should be a means to give unmistakable warning to the pilot when the lock is engaged. Several accidents have occurred because the pilot did not remove the control system lock prior to takeoff. Many such accidents relate to internally applied locks, mostly pins installed at the control wheel column. Misuse and alteration of these installed locking devices, together with neglect by the pilot to perform a control freedom check before takeoff, contributed to such accidents. When evaluating a control lock system, the following factors should be considered in finding compliance with the applicable regulation: a. The warning should be easily observable during both day and night operations. Color, location, shape, and accessibility of the device, ease of removal with the pilot seated in the flying position, and legibility of any placards, etc., should be considered. b. The system operation should be obvious. It should be possible to apply the lock only in such a manner that the required warning is provided. c. When engaged, the lock should, by design, limit the operation of the airplane so that the pilot receives unmistakable warning in the cockpit before or at the start of takeoff by an effective means, such as one of the following: (1) Preventing the application of sufficient engine power to attempt a takeoff. (2) Displacement of primary pilot controls, such as the control wheel full forward. (3) An aural warning device that cannot be disengaged. For airplanes with separate locks for throttle and control column, where one lock (e.g., throttle) can be removed independently of the other, each lock should independently meet the criteria of paragraph (c) above. ### Page 18 23.681 Limit load static tests No policy available as of June 30, 2001. ### Page 19 23.683 Operation tests Original Issue and Subsequent The 1.25 factor of part 23, § 23.395(a)(1) does not apply to the control system operational test of this section. Compliance with this section is required whether or not the airplane has a significant flight test history. Proof of structure is accomplished by ground tests because required flight tests may not subject the airplane to limit loads for all possible flight conditions. Amendment 23-7 and Subsequent Part 23, § 23.683, and CAR Section 3.343 require showing by operation tests, when the controls are operated from the pilot compartment with the system loaded, that the system is free from jamming, excessive friction, and excessive deflection. This section has not been uniformly applied. Some airplanes were certified using 50 percent of the control surface travel with no load as criteria for meeting the excessive deflection requirements for the operation tests. Other airplanes were not required to meet any specific travel as long as the airplane had adequate flight characteristics. Requiring a specific large travel while under limit load could result in control system authority that is greater than desired or needed. However, some travel of the control surface should exist when the system is loaded to limit load. No travel could indicate there was a possible fault, such as a jammed system. Secondly, with little or no travel, operation of the controls would have such a limited effect on the maneuverability of the airplane that it could have questionable flight characteristics. ACCEPTABLE MEANS OF COMPLIANCE One method, but not the only method, for showing compliance with the control system operation test requirements of § 23.683 and CAR Section 3.343 is as follows: a. This method may be used when clearances around control surfaces are no less than 3/16 inch. b. Conduct the control system operation tests by operating the controls from the pilot's compartment with the entire system loaded so as to correspond to the limit control forces established by the regulations for the control system being tested. The following conditions should be met: ### Page 20 (1) Under limit load, check each control surface for travel and detail parts for deflection. This may be accomplished as follows: (a) Support the control surface being tested while positioned at the neutral position. (b) Load the surface using loads corresponding to the limit control forces established in the regulations. (c) Load the pilot's control until the control surface is just off the support. (d) Determine the available travel, which is the amount of movement of the surface from neutral when the control is moved to the system stop. (e) The above procedure should be repeated in the opposite direction. (f) Minimum control surface travel from the neutral position in each direction being measured should be 10 percent of the control surface travel with no load on the surface. Regardless of the amount of travel of the surface when under limit load, the airplane should have adequate flight characteristics, as specified in § 23.141. Any derivative airplane of a previous type certificated airplane need not exceed the control surface travel of the original airplane; however, the flight characteristics should be flight tested to ensure compliance. (2) Under limit load, no signs of jamming or of any permanent set of any connection, bracket, attachment, etc., may be present. (3) Friction should be minimized so that the limit control forces and torque’s specified by the regulations may be met. ALTERNATE MEANS OF COMPLIANCE Applicants and FAA Aircraft Certification Offices (ACO's) involved with certification of small airplanes should generally follow this policy. Applicants should expect that the ACO would consider this information when making findings of compliance. However, in determining compliance with certification standards, each ACO has the discretion to coordinate deviations from these guidelines with the Small Airplane Directorate when the applicant demonstrates a suitable need. To ensure standardization, the ACO should coordinate deviation from this policy with the Small Airplane Directorate. Recently, airplanes have been built with smaller gaps between control surfaces and structure than has been done in the past. In this case, this alternate means of compliance is appropriate. This ### Page 21 11 method also can have the added advantage of demonstrating compliance with several regulations in a single test series. Regardless of the amount of travel of a control surface when tested as described above, the airplane must have adequate flight characteristics as specified in § 23.141. Any airplane that is a close derivative of a previous type certificated airplane need not exceed the control surface travel of the original airplane; however, the flight characteristics should be tested to ensure compliance. The method of showing compliance with § 23.683 presented in AC 23-17, paragraph 23.683, Operation Tests, discusses only the control system. It does not explicitly specify the consideration of loading on adjacent structures and elements. This is consistent with the wording in § 23.683 of the regulations. Testing, not analysis, must be used to show compliance with § 23.683. There are five other regulations, the control system, the control surfaces, and the adjacent fixed aerodynamic surfaces related to both the control system and the control surfaces, which must also be met. These include the following: 1. The first one is section 23.305, paragraph (a), [Subpart C - Structure, General] Strength and Deformation. It requires that "At any load up to limit loads, the deformation may not interfere with safe operation." 2. Section 23.307, [Subpart C - Structure, General] Proof of Structure, states that "Compliance with the strength and deformation requirements of § 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." 3. Section 23.655, paragraph (a), [Subpart D - Design and Construction, Control Surfaces] Installation, requires that "Moveable surfaces must be installed so that there is no interference between surfaces, their bracing, or adjacent fixed structure, when one surface is held in its most critical clearance positions and the others are operated through their full movement." 4. Section 23.681, paragraph (a), [Subpart D - Design and Construction, Control Surfaces] Limit Load Static Tests, requires that "Compliance with the limit load requirements of this part must be shown by tests in which — (1) The direction of the test loads produces the most severe loading in the control system; and (2) Each fitting, pulley, and bracket used in attaching the system to the main structure is included." ### Page 22 5. Section 23.141, [Subpart B--Flight, Flight Characteristics] General, states that "The airplane must meet the requirements of §§ 23.143 through 23.253 at all practical loading conditions and operating altitudes for which certification has been requested, not exceeding the maximum operating altitude established under § 23.1527, and without requiring exceptional piloting skill, alertness, or strength." To ensure that these requirements will be satisfied in the conduct of the control system operation test, inclusion of loads on the adjacent structures or elements in the testing set-up may be required. While testing is required for demonstration of compliance to § 23.683, in some cases analysis may be acceptable for showing compliance with § 23.305, paragraph (a). Section 23.307, paragraph (a), provides the criterion for when analysis is not acceptable and testing must be performed. It is not appropriate to define specific quantitative criterion to determine when testing is required to demonstrate compliance with § 23.305, paragraph (a), in accordance with 23.307, paragraph (a). One specific criterion will not work for all possible airplane designs. It is better that such determinations are made on a caseby-case basis, in which the appropriate details of a particular design can be considered. However, this policy describes some of the factors that should be considered when determining if tests are required to demonstrate that clearance between controls and adjacent structure (under load) meets 23.305, paragraph (a). These factors include, but are not limited to, the following: (1) The clearance between control surfaces and adjacent structure, when at rest. Suppose an applicant has experience with other airplanes that have a half-inch of clearance between controls and adjacent structure at rest. However, a new design is similar except it now has only a tenth of an inch clearance when at rest. Tests to demonstrate compliance with § 23.305, paragraph (a), may be required because the new structure may not conform to those for which experience has shown this method to be reliable in the past. The accuracy of past methods may not be suitable for the smaller clearances. Critical conditions assessed in past analysis may not have included a condition that is critical for the new smaller clearance. ### Page 23 (2) The amount of deformation (under limit loads) in the control surface or adjacent structure. If analysis had been shown to be reliable in the past for a wing that had much smaller deflections than a current design, the current structure may not conform to those for which experience has shown this method to be reliable, and testing may be required. Previous analytical methods may no longer be reliable because the new design behaves in a more nonlinear manner. It is possible that types of deflection that were neglected in past analysis may now become critical. (3) New control surface attachment configurations or other local design changes could create new types of deformation that are critical for the new design but were not considered in past analysis. If the FAA requires (or if an applicant voluntarily chooses) compliance with § 23.305, paragraph (a), to be shown by tests, the following test procedure is one means to simultaneously demonstrate compliance with both § 23.305, paragraph (a), and § 23.683. It also demonstrates compliance with § 23.681, paragraph (a). These tests may be conducted as follows: Except where otherwise specified, the tests described below in sections (1), (2), and (3) should be conducted within the following parameters (a through h). PARAMETERS: a. Conduct the control system operation tests by operating the controls from the pilot's compartment. b. All the control surfaces must be installed to their adjacent fixed surface on the airframe (according to the type design). c. The entire control system and adjacent fixed structure should be loaded. d. The adjacent fixed surfaces (wings, horizontal stabilizers, vertical stabilizers, and so forth) should be loaded to provide deflections equivalent to critical limit load flight conditions. e. The structural deflections should correspond to the limit flight conditions that represent the worst case conditions for increased ### Page 24 14 cable tension, decreased cable tension, and control/fixed surface proximity for each control system as appropriate. f. The entire control system must be loaded to either the limit airloads or the limit pilot forces, whichever is less (§ 23.683, paragraph (b)(1)). Per § 23.397, the automatic pilot effort must be used instead of limit pilot forces if it alone can produce higher control surface loads than the human pilot. g. Minimum clearances around control surfaces and minimum tensions in cable systems should be defined and incorporated in the airplane’s instructions for continued airworthiness. The test article should incorporate these minimum clearances and tensions, unless you otherwise account for them. h. If reductions in the minimum clearances described in paragraph g above are possible due to environmental conditions expected in service, you must account for this. This can be accomplished through analysis or during testing by adjusting the test article clearances to encompass these effects. SECTION (1): Consider all airplane maneuver and gust loads, and inertial loads, represented by the airplane flight envelope (V-n diagram); consider unsymmetrical load cases. (1) The tests described in this section support the demonstration that the control system is free from jamming, excessive friction, and excessive deflection as required by § 23.683, paragraphs (a)(1), (2), and (3). They also support the demonstration that structural deformations not interfere with safe operation as required by § 23.305, paragraph (a). Accomplish the following: (i) Load the adjacent fixed aerodynamic surface (wing, horizontal tail, or vertical tail) in accordance with one of the conditions of paragraphs d, e, and f above. (ii) Support the control surface being tested while it is located in the neutral position. (iii) Load the control surfaces to the critical limit loads, as described in paragraph f above, and evaluate their proximity to the fixed adjacent structure for interference (contact). ### Page 25 (iv) Load the pilot’s control until the control surface is just off the support. (v) Determine the available control surface travel, which is the amount of movement of the surface from neutral when the cockpit control is moved through the limits of its travel. (vi) The control surface under loads described in paragraph f above must have adequate flight characteristics as specified in § 23.141. (vii) To address the possibility of a critical intermediate control surface loading, gradually remove load from the control surface (while maintaining the load on the adjacent fixed surface) until maximum control surface travel is achieved. (viii) The above procedure should be repeated in the opposite direction. (ix) With limit load applied to the adjacent fixed surface and limit or intermediate load applied to the control surface, no signs of jamming, or of any permanent set of any connection, bracket, attachment, and so forth, may be present. (x) The control system should operate freely without excessive friction. Excessive friction is any increase under limit loads that results in exceeding the limit control forces and torques specified by the regulations. (xi) Cable systems should be checked with the loads applied to ensure that excessive slack does not develop in the system. Excessive slack is any change in cables or cable hardware that results in reduced airplane control surface movement. (xii) Repeat this process for each of the critical loading conditions as defined by paragraphs d and f above. SECTION (2): (2) The tests described in this section support the demonstration that structural deformations not interfere with safe operation as required by § 23.305, paragraph (a). Accomplish the following: (i) Load the adjacent fixed aerodynamic surface (wing, horizontal tail, or vertical tail) in accordance with one of the conditions of paragraph d and e above. (ii) Operate the unloaded control system from stop to stop. (iii) No signs of interference (contact) may be present. ### Page 26 (iv) The control system should operate freely without excessive friction. (v) Repeat this process for each of the critical adjacent fixed surface loading conditions as defined by paragraphs d and e above. NOTE 1: An alternate procedure may be used to accommodate the testing described in sections (1) and (2) above during structural tests of a partial airplane. This method requires that all control system components that are attached to or enclosed by the loaded test structure be installed per type design. A sufficiently representative mockup of remaining control system components must be used to ensure that the full length of any cables which extend from the loaded test structure are included. This is necessary to make a reasonable assessment that slack that could develop in control cables is not excessive enough to cause an entanglement or jam. The control surface activation may be input at any convenient location between the mockup terminus and the cockpit. SECTION (3): (3) The tests described in this section will demonstrate that the control system is free from excessive deflection as required by § 23.683, paragraph (a)(3). These tests complete this means of compliance that the control system is free from jamming and excessive friction as required by § 23.683, paragraphs (a)(1) and (2). They also demonstrate that structural deformations do not interfere with safe operation as required by § 23.305, paragraph (a). These tests meet the limit load static test requirements of § 23.681, paragraph (a). Accomplish the following: (i) With the adjacent fixed surface (wing, horizontal tail, or vertical tail) unloaded, support the control surface being tested while it is located in the neutral position. (ii) Load the control surfaces to the critical limit loads, as described in paragraph f above, and evaluate their proximity to the fixed adjacent structure for jamming or contact. (iii) Load the pilot's control until the control surface is just off the support. (iv) Operate the cockpit control in the direction opposite the load to the extent of its travel. (v) The above procedure should be repeated in the opposite direction. ### Page 27 (vi) The minimum loaded control surface travel must have adequate flight characteristics as specified in § 23.141. (vii) Under limit load, no signs of jamming, or of any permanent set of any connection, bracket, attachment, and so forth, may be present. (viii) The control system should operate freely without excessive friction. NOTE 2: The tests described in section (3) above are normally accomplished using a complete airplane. As a minimum, they must be completed using an airframe/control system that completely represents the final product from the cockpit controls to the control surface. Regardless of the amount of travel of a control surface when tested as described above, the airplane must have adequate flight characteristics as specified in § 23.141. Any airplane that is a close derivative of a previous type certificated airplane need not exceed the control surface travel of the original airplane; however, the flight characteristics should be tested to ensure compliance. JAA ACJ 23.683 is acceptable for FAA certification. ### Page 28 23.685 Control system details No policy available as of June 30, 2001. ### Page 29 23.687 Spring devices Compliance may be shown by flight tests with the spring disconnected to demonstrate that failure of the spring will not cause flutter or unsafe flight characteristics, or by performing a reliability analysis. The flight test option is preferred since this approach conclusively addresses the safety issue and is historically a minimal test burden. An applicant who chooses the reliability analysis option must accept that a failure of the spring could create unsafe flight characteristics and, therefore, meet the corresponding level of reliability. This may involve the use of redundant design such as dual springs and demonstration of flight characteristics with one removed. A positive determination of spring reliability requires that an applicant show the spring will perform its intended function for a specified interval under operational and environmental conditions appropriate for the proposed airplane. Although not normally used for structural substantiation, the reliability assessment methods for § 23.1309 may contain some concepts that would be helpful in demonstrating the reliability of the spring device. The reliability assessment should consider, but not necessarily be limited to, fatigue failures, failures due to corrosive environments, and any in-service changes in the spring characteristics, particularly the spring constant. The reliability assessment of a spring device used in any airplane flight control system must consider airworthiness standards other than § 23.687. These include, but are not limited to, flutter characteristics and handling qualities. In addition, § 23.687 is explicit in that the applicant must demonstrate reliability of the spring with tests that simulate service conditions. Tests are the required substantiation method; however, an applicant may show compliance with tests supported by analysis. In addition to the spring device testing requirement, 14 CFR part 23, § 23.601, General, requires testing for any design detail or part that has an important bearing on safety of flight. An applicant should also show compliance with 14 CFR part 23, § 23.305 (a) and (b). Finally, when a single spring is required for flutter, handling qualities, or any other regulatory reason, the spring should be considered a single path critical structure and meet the A-Basis requirements of § 23.613(b)(1). ### Page 30 23.689 Cable systems Original Issue and Subsequent If tabs are installed with cable less than 1/8th inch diameter, the airplane should be safely controllable with the tabs in the most adverse position as if from a failed cable. Using emergency procedures, the pilot should be able to return and land safely. Airplane configurations, such as flaps, landing gear, and power are permissible devices to use in relieving control forces. The temporary control forces of part 23, § 23.143, are applicable until the force reduction procedures are completed. Smaller diameter cables (no less than 1/16 inch diameter) may be used for rudder pedal interconnections (used for pulling one pedal back when the other is pushed forward, but not used to drive the control surface), if the failure of this interconnection will not affect rudder operation. This rule was put in the original issue of part 23 from CAR 3.345, and it was intended to apply to airframe control cables not engine controls cables, which are certified under part 23, Subpart E. ### Page 31 23.691 Artificial stall barrier system Amendment 23-49 and Subsequent Section 23.201(b), Amendment 23-45, added the activation of an artificial stall barrier as an acceptable means of identifying when a stall has occurred. A stall barrier is a device that prevents an actual stall (i.e., a stick pusher) while a stall warning is a device that alerts a pilot of an impending stall (i.e., a stick shaker). Of course, the actual stall should not occur before activation of the stall barrier. This amendment provided the standards for an artificial stall barrier system when it is used to show compliance with § 23.201(b). Per § 23.201, a stick shaker is a "stall avoidance" device. Per the NPRM for Amendment 23-45, an applicant may identify a stall as the speed at which a stick shaker activates. This is an acceptable means of compliance to the rules. Stick shaker activation is then identified as a stall for a pilot by AFM. However, from an engineering perspective and considering the "stall avoidance" in § 23.201, this is an acceptable means of compliance not an aerodynamic stall. The difference is important in certification since the aerodynamic stall determines whether a stick shaker is required equipment. A stick pusher system would be a critical system for an airplane with stall recovery that is undetermined, marginal, or unacceptable. Failure of the system is then required to be extremely improbable. The FAA does not consider the probability of entering a stall environment as a factor in developing system reliability. The exception would be developing specific system component reliability where that component would be active only when the airplane is in a stall environment. The FAA does not give credit toward developing reliability for the use of a ‘Go/No Go’ preflight system check, although the FAA does recommend that preflight procedures for all essential/critical systems be provided for pilot use. (Service experience has shown that some part 23 airplane pilots do not have the discipline to conduct the prescribed preflight checks.) The development of normal/abnormal/emergency procedures is not a factor in determining system reliability; however, such procedures are desirable, as well as required by § 23.1581. These factors may be considered when exercising engineering judgment in approval of the overall system. Stall may be identified by stick shaker/pusher operation, uncontrollable downward pitching, or the elevator control reaching the stop (see AC 23-8A)—whichever occurs first in any particular flight regime is acceptable. An airplane may be approved if it has stick shaker/pusher operation in one configuration, such as power on, and it has acceptable stall characteristics for the remaining configurations. Inadvertent stick pusher operation should be investigated and shown not to be hazardous and to be recoverable, or that inadvertent operation is extremely improbable. ### Page 32 ### Page 33 23.693 Joints No policy available as of June 30, 2001. ### Page 34 23.697 Wing flap controls No policy available as of June 30, 2001. ### Page 35 23.699 Wing flap position indicator No policy available as of June 30, 2001. ### Page 36 23.701 Flap interconnection Original Issue The flaps should be synchronized by a mechanical interconnection unless the airplane has safe flight characteristics with the flaps retracted on one side and fully extended on the other side. The safe flight demonstration with asymmetry should be shown throughout the airspeed range permitted for flap extension. The control forces should not exceed those shown for temporary application in the table in § 23.143(c). However, they may not exceed the force that can be demonstrated as safe with one hand on the control wheel/stick (other hand needed to re-trim, pull circuit breaker, operate flap control, etc.). If the forces of asymmetry cannot be alleviated in a reasonable period of time, the remaining forces should not exceed those specified for prolonged application in § 23.143(c). After demonstrating that the airplane has safe flight characteristics with the flaps in their most adverse position, it is permissible to readjust the remaining flap surfaces after a malfunction occurs. Amendment 23-42 Amendment 23-42 was not intended to change the requirement that “The main wing flaps and related movable surfaces as a system must be synchronized by mechanical connection. . . .” The main purpose of this change was to add the following requirement that would maintain synchronization so that the occurrence of an unsafe condition has been shown to be extremely improbable. This requirement includes provisions for synchronization of the flaps other than by mechanical interconnection of the flap. These reliability requirements by numerical probability analysis for other synchronization methods should not be applied to mechanical interconnection. It is difficult to assess the reliability of mechanical interconnections by examples of different types of mechanisms. The complete system needs to be analyzed and tested. Section 23.701, as amended by Amendment 23-42, in part, states the following: (a) The main wing flaps and related movable surfaces as a system must: (1) Be synchronized by mechanical connection; or (2) Maintain synchronization so the occurrence of an unsafe condition has been shown to be extremely improbable; or ### Page 37 (b) The airplane should be shown to have safe flight characteristics with any combination of extreme positions of individual movable surfaces (mechanically interconnected surfaces are to be considered as a single surface). During a recent review of this new requirement, it was noted that the new § 23.701(b), particularly the parenthetical portion of that paragraph, could be improperly interpreted and applied. It is possible that this misinterpretation could result in the use of differing terminology (i.e., "mechanical interconnection" and "mechanically interconnected") in paragraphs (a)(1) and (b). These terms mean the same thing; direct positive mechanical interconnection between separate flap surfaces that are isolated from the flap control or actuation system. Novel and unusual design features, such as an interconnection of the leading and trailing edge flap systems or an interconnection of flaps and ailerons, would require special conditions. Equivalent Level of Safety Findings Several findings have been accepted for the mechanical interconnection requirement. a. Synchronized by a Mechanical Interconnection These words appeared in 14 CFR parts 23 and 25 and in CAR 03 and 04b since they were first issued. The synchronization requirement for the motion of the flaps by a mechanical interconnection is applicable to airplanes not having safe flight characteristics under asymmetrical flap operations. For these cases, there would be a hazardous condition when the flaps are retracted on one side and extended on the other side. b. Mechanical Interconnection Requirement of § 23.70l(a)(l) This requirement is to ensure against hazardous asymmetrical operation of the flaps after any probable single or probable combination of failures of the flap actuating system. A probable combination of failures should be considered when the first failure would not be detected during normal operation of the system, including periodic checks, or when the first failure would inevitably lead to other failures. (Systems where a probable combination of failures may occur include the electrical and hydraulic systems.) The airplane also should be shown to be capable of continued safe flight and landing without requiring exceptional pilot skill or strength following these failures. To demonstrate that the airplane is safe under these conditions, tests should be conducted with the flaps being retracted on one side and extended on the other during takeoffs, approaches, and landing. If there is a probable hazardous condition, a separate positive connection that is not part of the flap actuation system is required. ### Page 38 ACCEPTABLE MEANS OF COMPLIANCE An acceptable means of compliance with the airworthiness requirements for the flap's mechanical interconnections of § 23.701(a)(1) are described below: a. Reliability Reliability of the mechanical interconnections is generally shown either by load analysis or load tests, or both, not by numerical probabilistic analysis. The mechanical interconnection should be designed for the loads resulting when interconnected flap surfaces on one side of the plane of symmetry are jammed and immovable, while the surfaces on the other side are free to move and the full power of the surface actuating system is applied. It should also be designed to account for the asymmetrical loads resulting from flight with the engines on one side of the plane of symmetry inoperative and the remaining engines at takeoff power. For single engine airplanes and multiengine airplanes with no slipstream effects on the flaps, it may be assumed that 100 percent of the critical air loads acts on one side and 70 percent on the other. The flight loads from § 23.345 acting on the surfaces should be considered in combination with the actuating system loads (including system inertia loads). Critical air load conditions should consider flap retraction and flap extension, including go-around. These conditions are considered limit loads. If there are no hazardous conditions when the flaps are asymmetrical, the jam or maximum load conditions could be considered an ultimate load. b. Friction Loads It may be necessary to consider friction loads in the actuating system that may be reasonably expected to occur in service. Each design should be evaluated to determine its susceptibility to friction in the mechanism and any loads with such resistance. c. Equivalent Means by Use of the Mechanical Actuation System The mechanical actuating system for the flaps may be considered the mechanical interconnection, if all elements are mechanically interconnected from the actuator source to the flaps. These mechanical elements may include structures, interconnection linkages, and drive system components. When the mechanical interconnection is through the actuating system, and it is the only means to prevent an unsafe asymmetrical condition, the loads associated with the jam conditions are considered limit loads. A 1.5 factor of safety is required if a failure as a result of the jam condition would cause a hazardous flap asymmetrical operation. A mechanical actuating system having a 1.5 factor of safety may not need to be evaluated for probable failure conditions. Also, if the drive system is designed so that a hazardous flap asymmetrical operation would not occur after a jam condition, the 1.5 factor of safety should not be required. ### Page 39 d. Equivalent Means by Use of a Warning and Prevention System A second equivalent means is the use of a warning and prevention system. This system monitors the symmetrical condition of the flaps and warns the pilot when an unsymmetrical flap condition occurs, but the asymmetry is still kept within safe limits. It prevents further movement of the flaps from exceeding safe limits. The warning and prevention system should be independent for each functionally related set of surfaces (i.e., a set of flaps on each side of a plane of symmetry that is driven by a common actuator). Again, the airplane should be shown to have safe flight characteristics without requiring exceptional piloting skill or strength at the extreme limits of the asymmetrical condition where the flaps are stopped. Tests should be conducted to simulate flap malfunctioning at the most severe case in the static asymmetrical condition of the flaps during takeoffs, approaches, and landings. The warning and prevention system should provide a pilot with a selectable or automatic test mode that exercises the system to an appropriate depth, so the pilot can determine proper operation of this system. e. Electrical/Electronic Flap Interconnection System When Amendment 23-42 was adopted, § 23.701 was amended to include provisions for airplanes with a flap configuration other than a mechanical interconnection. This amendment added the following requirement in § 23.701(a)(2): “Maintain synchronization so that the occurrence of an unsafe condition has been shown to be extremely improbable." This requirement is applicable for electrical/electronic flap interconnection systems, such as airplanes that have additional flaps and tandem wings. Guidelines for performing a design safety assessment by application of § 23.1309(b), as adopted by Amendment 23-41, are given in AC 23.1309-lC. This AC also provides guidance regarding design safety assessments, environmental and atmospheric conditions, and software assessment. ### Page 40 23.703 Takeoff warning system No policy available as of June 30, 2001. ### Page 41 **LANDING GEAR** 23.721 General The Small Airplane Directorate has no criteria or policy for taxi over rough surfaces for small airplane landing gear design. Manufacturers are free to develop and substantiate their own criteria. ### Page 42 23.723 Shock absorption tests Original Issue This regulation requires shock absorption tests be performed for certification. Amendment 23-23 and Subsequent This amendment permits an analysis rather than a shock absorption test, but to do so the applicant should have a landing gear system with identical (similar is not acceptable) energy absorption characteristics. The energy absorption characteristics of the landing gear system (e.g., structure, wheel tire, shock absorber) should be included in determining the dynamic response of the landing gear system. The tests should cover a range of energy absorption characteristics and weights over which the analysis is shown to be valid. If these conditions are not met, drop tests will be required to substantiate maximum takeoff and landing weight increases. It is acceptable to modify individual gear drop test data by adapting the results to the complete aircraft analytically, accounting for the aircraft flexibility. ### Page 43 23.725 Limit drop tests Original Issue This rule gives requirements for limit load drop tests if the applicant uses free drop tests to meet the requirements of § 23.723(a). The applicant should make ten drops from limit height for each basic design condition. The applicant should make one drop from the height (maximum is 2.25 times the limit drop height) needed to develop 1.5 times the limit load using the limit drop weight. It is acceptable to modify individual gear drop test data by adapting the results to the complete aircraft analytically, accounting for the aircraft flexibility. Amendment 23-7 and Subsequent This amendment requires that the limit inertial load factor be determined in a rational and conservative manner during the drop test using a landing gear unit attitude and applied drag loads that represent the landing conditions. It is acceptable to modify individual gear drop test data by adapting the results to the complete aircraft analytically, accounting for the aircraft flexibility. ### Page 44 23.726 Ground load dynamic tests No policy available as of June 30, 2001. ### Page 45 23.727 Reserve energy absorption drop test Amendment 23-7 and Subsequent Paragraph (b) in § 23.727 requires that the effect of wing lift be provided for in reserve energy drop tests. You should also use the applicable drag loads, as specified in § 23.725(c). ### Page 46 23.729 Landing gear extension and retraction system Original Issue No policy available as of June 30, 2001. Amendment 23-7 and Subsequent A warning device with no manual shutoff is required when the flaps are “to or beyond” the approach flap setting if the landing gear is not down and locked. The “to or beyond” phrase in relation to using a normal landing procedure is intended to provide for differences in design, as follows: a. For airplanes whose normal procedures only prescribe landings with flaps extended past the approach setting, only the “beyond” aspect of this rule is appropriate. Operating information for these airplanes should convey that landings with approach flaps, or less, are not normal and will not activate the flap/landing gear aural warning. b. For airplanes whose normal procedures include landings with a flap setting at the approach setting, the “to and beyond” aspect of this rule is appropriate. Designers may choose to include additional logic in the flap/landing gear warning system, such as airspeed, thrust/throttle position, etc. This logic may tend to minimize nuisance warnings and may provide the equivalent safety intended by the rule. Because part 23 is not specific with regard to flap positions used, we cannot specify the flap position that actuates the warning device. This rule provides a basis for the FAA and the applicant to mutually agree on the set point for the warning device. Although not defined in part 23, most airplanes do have a “normal landing procedure” and an “approach flap position.” The flap position will vary among models, but it is this position that should be used to show compliance. This information is applicable to the structural substantiation to the loads resulting only from all yawing conditions for the landing gear doors and retraction mechanism of small airplanes per part 23, § 23.729(a)(2). Section 23.729(a)(2) requires the landing gear doors and retraction mechanism to be substantiated for the loads resulting from all yawing conditions. Attempts have been made to meet these requirements by flight testing to dive speed with some yaw or by flight testing at full yaw at a lower speed. These procedures normally do not result in ### Page 47 37 a test that substantiates a 1.5 factor of safety. If substantiation by flight testing is desired, the landing gear doors and retraction mechanism should be subjected to 1.5 times the limit "q" loading. The limit “q” loading is the “q” at VLE or VLO, whichever is greater. The higher of the above speeds at which Vq is to be computed is designated as VLG. ACCEPTABLE MEANS OF COMPLIANCE One method, but not the only method, for showing compliance with the structural requirements of § 23.729(a)(2) for the loads resulting from all yawing conditions for the landing gear doors and retraction mechanism is as follows: a. Substantiation may be accomplished by flight testing at a speed of Vq and the yaw angle determined in paragraph a(3) below, unless this will exceed the structural limitation as determined by analysis, static test, or a combination of both, where: (1) Vq = square root of (VLG squared times 1.5). (2) VLG = The greater of VLO or VLE. (3) For the yawed condition, the limit “q” load will be at VLG with the airplane at the yaw angle determined by § 23.441. This angle need not exceed 15o. Substantiation should be to 1.5 “q”. (4) If Vq is equal to or less than VA, substantiation by flight test may be accomplished. (5) If Vq is greater than VA, the yaw necessary to produce 1.5 “q” could result in overloading other airplane structures, and the maneuver should not be performed. (6) VLG may be reduced by imposing limitations on the airplane such that Vq is less than VA. ### Page 48 (7) The definitions of the terms used above are equivalent airspeeds, as follows: VA = Design maneuvering speed VD = Design diving speed VLG = Landing gear speed used in the calculation of Vq VLE = Maximum landing gear extended speed VLO = Maximum landing gear operating speed Vq = Speed which results in 1.5 times limit "q" loading b. If the condition of Item a.(5) above exists, substantiation of the landing gear doors and retraction mechanism may be accomplished by static tests, analyses, or a combination of both. Amendment 23-49 and Subsequent This guidance provides clarification of the requirements for tire burst as related to landing gear and 14 CFR part 23, §§ 23.729(g) and 23.1309 compliance. The requirement in § 23.729(g) states: “If the landing gear bay is used as the location for equipment other than the landing gear….” As used within this requirement, the term landing gear applies to all parts and systems contained in the “LANDING GEAR” section in Subpart D (§§ 23.721 through 23.745) that are specific to the landing gear. For example, wheels, brakes, wheel steering mechanism, and structural portions of the landing gear that are within the gear are all considered part of the “landing gear” for the purposes of § 23.729(g) compliance. Systems that act upon the landing gear, e.g., extension and retraction systems, hydraulic systems, etc., are not considered part of the landing gear itself and should be addressed in § 23.729(g) compliance. Compliance to this section can be by isolation of non-landing gear equipment from the landing gear bay or by protection/shielding of this equipment from the effects of tire burst/loose tread and external forces. Using either of these means of compliance to § 23.729(g) also shows compliance to the § 23.1309 Particular Risk Analysis for non-landing gear equipment. The effects of tire burst must still be evaluated for § 23.1309 compliance. When showing compliance with § 23.1309 for tire burst, in accordance with § 23.1309(f)(2) and Advisory Circular 23.1309-1C, only the structural portions of the landing gear are exempt from the § 23.1309 requirements. Addressing external effects as part of the ### Page 49 39 airplane safety analysis is consistent with the FAA advisory material for parts 23, 25, and 29, Aerospace Recommended Practice (ARP) 4761, and “Guidelines and Methods for Conducting the Safety Assessment Process on Civil Airborne Systems and Equipment.” When showing compliance with § 23.1309 for tire burst, it is not appropriate to address compliance by assuming a probability of tire burst occurring, rather tire burst must be assumed to have occurred(assuming a tire burst probability is inconsistent with the aforementioned advisory material and accepted industry guidelines). Therefore, an acceptable means of compliance with the tire burst requirements § 23.1309 would be to account for all systems and equipment in the tire burst tests and analysis performed under § 23.729(g). These tests and analyses must be conducted on a conformed test article that includes all systems and components as defined in the type design data. Items identified by either analysis or tire burst tests, or both, as capable of causing hazards to the airplane must be protected or relocated. Therefore, we provide the following as answers to the questions raised: 1. “Do hydraulic, mechanical, and electrical components utilized in any way for retraction/extension of the landing gear doors or control of same need to be protected from the tire burst requirements of § 23.729(g)?” Yes, unless the § 23.1309 safety assessment shows that damage from tire burst cannot cause a hazard(s) to the airplane. The minimization requirements of § 23.729(g) apply. 2. “If the above items are exempt from § 23.729(g), are they exempt from the Particular Risk Analysis required by § 23.1309.” While § 23.729(g) applies to the Item 1 equipment, tire burst effects should be considered for all landing gear and non-landing gear systems and equipment, except flight structure, as part of § 23.1309 compliance. ### Page 50 23.731 Wheels See 23.735, Brakes. ### Page 51 23.733 Tires Original Issue and Subsequent The following is a recommended test procedure for the installation of tires on a part 23 airplane: 1. Inflate an inboard main tire to the minimum allowable inflation pressure for the airplane weight. 2. Inflate the outboard main tire on that same landing gear to the maximum allowable inflation pressure for the airplane weight. 3. Using white shoe polish or equivalent, mark a 2-inch wide stripe on the brake (inboard) side of the outboard tire sidewall adjacent to the wheel rim. 4. Conduct at least two maximum effort, non-skidding taxi turns into the minimum inflation side of the airplane. 5. Check for evidence of brake wheel housing abrasion contact on the tire sidewall. Note: Above applies to a dual tire installation per landing gear. For a single tire per gear, inflate either side to the minimum pressure and the opposite side to maximum, and turn into the minimum pressure side. Tundra Tires 1. PURPOSE. This guidance serves several purposes. First, it summarizes the results of flight tests recommended by the National Transportation Safety Board (NTSB) and conducted by the Federal Aviation Administration (FAA) to investigate the effects of tundra tires installed on a Piper PA-18. Second, it provides information concerning possible hazards associated with the type of operations common for tundra tire users and potential adverse effects of untested installations. Third, it provides general information about the certification process for oversized "tundra" tires, as well as an example "compliance checklist" for the installation of such tires on light airplanes that have Civil Air Regulations (CAR) part 3 for a certification basis. **2. RELATED READING MATERIAL** a. Part 23, CAR part 3, and CAR part 03. b. National Transportation Safety Board (NTSB) Safety Recommendation A-95-13, dated February 7, 1995. ### Page 52 c. Technical Standard Order (TSO)-C62d, Aircraft Tires. d. AC 43.13-1B, Acceptable Methods, Techniques, and Practices—Aircraft Inspection and Repair. 3. BACKGROUND. In Safety Recommendation A-95-13, dated February 7, 1995, the NTSB shared some of their safety concerns about tundra tires with the FAA and requested that the possibility of problems with tundra tires be investigated. The NTSB stated the following: “Since the early 1960’s, hundreds of airplanes operating in Alaska have been equipped with tundra tires, and dozens of versions of tundra tires—some exceeding 35 inches in diameter—have been marketed. The Safety Board is concerned that field approvals and STC’s have been granted for use of these tires without flight test or other data on the aerodynamic effects of the tires and wheels. The Piper PA-18 is the airplane most frequently equipped with tundra tires. The Safety Board believes that the FAA should conduct a demonstration flight test to determine the effects of tundra tires on the PA-18's flight characteristics—including cruise, climb, takeoff, and landing performance—and, in both straight and turning flight, stall warning and aircraft stability at or near the critical angle of attack. Further, if the tests of the PA-18 indicate the need, the FAA should take corrective action and expand testing to other airplane types equipped with oversized tires.” **4. SUMMARY OF FLIGHT TEST RESULTS FOR PIPER PA-18 EQUIPPED** WITH TUNDRA TIRES The FAA's flight tests of tundra tires and their results are detailed in Appendix 1 following this guidance. As can be seen in the report, the tundra tire installations on the Piper PA-18 “150” caused no observable adverse effects on stall or stall characteristics during the FAA tests. Although there was some degradation of handling qualities associated with increasing the tire size, the effect was not significant with regard to safety. Rate of climb and cruise speed were degraded with the larger tire sizes; however, the aircraft still met certification requirements. Additional tests conducted by an independent Designated Engineering Representative (DER) flight test pilot showed the same lack of effect on stall characteristics with the main landing gear fabric covering removed. It should be remembered that these results are valid only for the Piper PA-18 “150” and for tires no larger than those tested. It should also be noted that, although tundra tires did not cause a safety problem, the stall characteristics of the basic Super Cub (and most other airplanes) make low altitude turning stalls hazardous, especially in uncoordinated flight. Also, although washout was not varied during these flight tests, previous FAA experience has shown that stall ### Page 53 43 characteristics are further aggravated when operators of the PA-18 remove the 2.5° of washout at the wing tip, which is not an approved alteration. This condition will result in a rapid roll when the airplane is stalled during turning flight. **5. POTENTIAL ADVERSE EFFECTS OF TUNDRA TIRE INSTALLATIONS** **ON AIRPLANES** a. Performance Tundra tire installations on airplanes may produce one or more of the following effects on performance characteristics: (1) Increased stall speed. (2) Reduced stall warning margin. (3) Reduced rate of climb. (4) Reduced maximum angle of climb. (5) Reduced maximum level flight speed. (6) Reduced cruise speed. (7) Reduced range. Tundra tires reduce climb, cruise, and range performance more when installed on relatively “clean,” well streamlined airplanes than they do when installed on less streamlined airplanes. b. Flight and Ground Handling Characteristics Tundra tire installations on airplanes may produce one or more of the following effects on handling characteristics: (1) Reduced ability of brakes to hold against takeoff power. (2) Reduced brake effectiveness during rejected takeoff and braked landing. (3) Reduced stability and controllability during rejected or balked landing and go around. (4) Change in either trim range or trim authority, or both. ### Page 54 (5) Reduced directional stability and control during takeoff and landing ground rolls, with consequent increased tendency to ground loop. (6) Increased tendency to nose over during landing. (7) Reduced stall warning margin, change in either aerodynamic stall warning characteristics (warning buffet) or reduced effectiveness of stall warning system, or both, in both level and turning flight with power either on or off, or both. (8) Changes in stalling and stall recovery behavior in both level and turning flight with power either on or off, or both. Stalls may become more abrupt and altitude loss before recovery may increase. (9) Increased tendency to enter an inadvertent spin and reduced ability to recover from the spin. (10) Reduced longitudinal, lateral, and directional stability. (11) Increased airframe vibration and buffet. Tundra tires reduce the airplane’s directional stability and controllability during takeoff and landing ground rolls, increase its tendency to ground loop during takeoff and landing ground rolls, and increase its tendency to nose over during landings on paved surfaces more than during landings on gravel, grass, or other surfaces that allow the tires to skid more easily. ### Page 55 c. Potential Propulsion Systems Effects (1) Fuel flow may be affected by changes in normal flying attitude. (2) Unusable fuel may be affected by changes in normal flying attitude. (3) The fuel tank sump may be affected by change in ground attitude. (4) Fuel drains may be affected by change in ground attitude. (5) Engine cooling may be affected by performance or flying attitude changes. (6) Changes in tires may affect the air induction system certification regarding operation on wet runways. (7) Changes in tires that change ground attitude may affect the induction system icing protection. **6. CERTIFICATION OF TUNDRA TIRES FOR USE ON LIGHT** AIRPLANES The certification process for tundra tires is the same as for any other tire to be used in aviation. a. A manufacturer may obtain a Technical Standard Order Authorization (TSOA) for the tire using the requirements in TSO-C62d. TSO-C62d contains minimum performance standards for aircraft tires. The TSOA, which covers design and manufacturing of the tire only, is not an installation approval. The tire should be approved for installation on a specific airplane model via Type Certificate (TC) or Supplemental Type Certificate (STC). The applicable requirements for installation of a tire on a given airplane should be determined based upon the original certification basis specified in that airplane’s Type Certificate Data Sheet. The development of a compliance checklist, as described in Item 7 below, should be accomplished by the applicant together with the FAA engineer. b. An alternative certification method exists for a tire that does not have a TSOA. In such a case, the tire design approval may be obtained concurrently with the installation approval for specific airplane models via TC or STC. The requirements of the TSO can be used for a determination of acceptable tire performance in such a project. The applicable requirements for installation of a tire on a given airplane should be determined based upon the original certification basis specified in that airplane’s Type Certificate Data Sheet. The development of a compliance checklist, as described in Item 7, should be accomplished by the applicant together ### Page 56 46 with the FAA engineer. Prior to offering tires approved by this method for sale, the tire manufacturer would need a Parts Manufacturing Approval (PMA). 7. COMPLIANCE CHECKLIST See Appendix 2 for an example of the "Compliance Checklist," to CAR part 3 as amended to November 1, 1949. This checklist is intended to show all aircraft certification requirements that could be affected by a tundra tire installation. Many of these requirements may be unaffected by a given installation. The actual compliance checklist for a specific installation should be determined at the start of a project. (See Appendix 1 and Appendix 2 in this section, which are applicable to § 23.733 Tundra Tires.) ### Page 57 47 This page intentionally left blank. ### Page 59 Appendix 1 47 APPENDIX 1 **FAA TEST RESULTS/EFFECTS OF TUNDRA TIRES** **ON THE HANDLING QUALITIES/STALLS/STALL CHARACTERISTICS OF THE PIPER PA-18** 1. Tests Recent accidents in Alaska involving airplanes equipped with tundra tires prompted the National Transportation Safety Board to recommend to the FAA that they conduct flight tests to determine the effects of tundra tires on aircraft performance, stalls, and handling qualities. The following five tires were evaluated at various combinations of center of gravity/weight: a. Factory installed (8.00-6) b. McCreary Tundra Tires (8.50-10) c. McCreary Tundra Tires (29x11.0-10) d. Schneider Racing Slicks (14.0x32.0x15) e. Goodyear Airwheels (35x15.0-6) 2. Results Quantitative/qualitative data obtained from the testing of the four tundra tires were compared to the data obtained from the testing of the factory-installed tire. The following is a summary of the findings: a. Ground Handling Forward field of view during taxi is inversely related to tire size. As the tire size increases, the ability to see over the nose decreases requiring that the pilot make “S” turns with the airplane. Ground handling during takeoff from a gravel runway is satisfactory for all configurations. Ground handling during landing on a gravel runway is also satisfactory for all configurations tested, although there is a noticeable nose down pitching moment when the tire(s) contact the ground. This is most evident when making a main wheel only landing. Crosswind landings on runway 13 at Lake Hood Strip, a 2,200′ x 80′ gravel runway next to Lake Hood 3 miles southwest of Anchorage, Alaska, were demonstrated for tire configurations Items 1a, 1b, and 1c in winds from 180° (from ahead and to the right of the airplane at an angle of 50° to its flight path) at 14 knots gusting to 16 knots. The wind thus had a crosswind ### Page 60 Appendix 1 48 component of approximately 10.7 knots gusting to approximately 12.3 knots and a head wind component of approximately 9.0 knots gusting to approximately 10.3 knots. No crosswinds were available during tests for configurations d and e. No tests for ground handling were accomplished on paved runways. The ground handling characteristics of airplanes equipped with tundra tires are known to be substantially poorer on pavement than on gravel, grass, and other surfaces that allow the tires to skid easily. b. Performance Tundra tires adversely affect airplane performance. For example, the uncorrected average rate of climb (tested at 1.05 times maximum gross weight) for the standard tire was 526 feet per minute. The uncorrected average rate of climb for configurations Items 1d and 1e (tested at 1.05 times maximum gross weight) was 449 and 464 feet per minute, respectively. 3. Stalls/Stall Characteristics a. The purpose of the stall tests was to determine whether there are any differences between the stalling speed and stall characteristics of a PA-18 ‘150’ airplane equipped with tundra tires and the stalling speed and stall characteristics of the same airplane equipped with standard tires. The data obtained from the stall tests do not validate the theory that tundra tires increase the PA-18 ‘150’ stalling speed. b. Stall characteristics (all configurations) are normal when the airplane is stalled in balanced flight. In a turning stall, the airplane generally rolls slowly to a near wings level attitude. In maneuvering flight, the tendency is for the nose to drop as the bank angle is increased. If the pilot uses top rudder (right rudder in a left turn) to compensate for this and then stalls the airplane, the airplane may roll rapidly over the top. This could result in a departure or the incipient phase of spin. If the airplane is maneuvering at low altitude when this sequence of events occurs (e.g., while circling to spot moose), the airplane may impact the ground prior to recovery. Also, although washout was not varied during these flight tests, previous FAA experience has shown that stall characteristics are further aggravated when operators of the PA- 18 remove the 2.5° of washout at the wing tip, which is not an approved alteration. This condition will result in a rapid roll when the airplane is stalled during turning flight. 4. Handling Qualities For any given center of gravity/weight, the lateral and directional stability tends to deteriorate as tire size is increased. ### Page 61 Appendix 1 49 5. Stall Warning Installation of the artificial Stall Warning System on the PA-18 is optional. Most of the PA-18's in Alaska do not have the system installed. The airplane tested did have the artificial Stall Warning System, and a number of test points were obtained with the system deactivated. The airplane as tested does not have an aerodynamic stall warning. ### Page 62 Appendix 2 50 APPENDIX 2 A "COMPLIANCE CHECKLIST" TO CAR PART 3, AS AMENDED TO NOVEMBER 1, 1949 Subpart B—Flight Requirements Weight Range and Center of Gravity Section Subject 3.71 Weight and balance 3.72 Use of ballast 3.73 Empty weight 3.74 Maximum weight 3.75 Minimum weight 3.76 Center of gravity position Performance Requirements—General 3.81 Performance* 3.82 Definition of stalling speeds* 3.83 Stalling speed* Takeoff 3.84 Takeoff* Climb 3.85 Climb* Landing 3.86 Landing* Flight Characteristics 3.105 Requirements* (exclude § 3.117) Ground and Water Characteristics 3.143 Requirements* ### Page 63 Appendix 2 51 Flutter and Vibration Section Subject 3.159 Flutter and vibration (vibration only) * Indicates topics identified by NTSB Safety Recommendation A-95-13 Subpart C—Strength Requirements Symmetrical Flight Conditions (Flaps Retracted) 3.189 Airplane equilibrium Flaps Extended Flight Conditions 3.190 Flaps extended flight conditions Unsymmetrical Flight Conditions 3.191 Unsymmetrical flight conditions Control Surface Loads 3.211 General 3.212 Pilot effort 3.213 Trim tab effects Horizontal Tail Surfaces 3.214 Horizontal tail surfaces Vertical Tail Surfaces 3.219 Maneuvering loads Control System Loads 3.231 Primary flight controls and systems Ground Loads 3.241 Ground loads 3.242 Design weight 3.243 Load factor for landing conditions ### Page 64 Appendix 2 52 Landing Cases and Attitudes Section Subject 3.244 Landing cases and attitudes Ground Roll Conditions 3.248 Braked roll 3.249 Side load Subpart D—Design and Construction Control Systems 3.342 Proof of strength Landing Gear 3.351 Tests 3.352 Shock absorption tests 3.353 Limit drop tests 3.354 Limit load factor determination 3.355 Reserve energy absorption drop tests Wheels and Tires 3.361 Wheels 3.362 Tires Brakes 3.363 Brakes Subpart E- Propulsion Systems 3.429 Fuel System General 3.437 Unusable Fuel 3.444 Fuel Tank Sump 3.553 Fuel Tank Drains 3.583 Cooling Tests 3.605 Induction System General 3.606 Induction Systems de-icing and anti-icing provisions ### Page 65 Appendix 2 53 Subpart F—EQUIPMENT Landing Lights 3.699 Landing light installation Subpart G—OPERATING LIMITATIONS AND INFORMATION 3.735 General Limitations 3.737 Limitations ### Page 66 23.735 Brakes Original Issue and Subsequent Brake Replacement and Modified Parts 1. Related regulations and documents are: a. Regulations Acceptable means of compliance are found in 14 CFR part 23. Additional specific information is listed below, including other regulatory material and advisory information. Part 23 sections may be used in showing compliance with the corresponding sections of the former Civil Air Regulations (CAR) for airplanes where the CAR regulations are applicable. For convenience, the former CAR section reference is also shown in parenthesis following the part 23 section reference: Part 21, § 21.15 Application for type certificate Part 21, § 21.93 Classification of changes in type design (TC) Part 21, § 21.113 Requirement of supplemental type certificate Part 21, § 21.303 Replacement and modification parts Part 21, § 21.611 Design changes (TSO) Part 23, § 23.55 Accelerate-stop distance Part 23, § 23.75 (3.86) Landing Part 23, § 23.143 (3.106) Controllability and maneuverability: General Part 23, § 23.231 (3.144) Longitudinal stability and control Part 23, § 23.233 (3.145) Directional stability and control Part 23, § 23.493 (3.248) Braked roll conditions ### Page 67 Appendix 2AC 23-17A 55 Part 23, § 23.731 (3.361) Wheels Part 23, § 23.735 (3.363) Brakes Part 23, § 23.1301 (3.652) Function and installations Part 23, § 23.1309 Equipment, systems, and installations Part 23, Appendix D Wheel spin-up and spring back loads Part 23, § 23.1529 and Instructions for Continued Airworthiness Appendix G to Part 23 Part 135, Appendix. A Additional airworthiness standards for 10 or more passenger airplanes Part 45, § 45.14 Identification of critical components Part 45, § 45.15 Replacement and modification parts b. Advisory Circulars (AC's) AC 21.303-2H Announcement of Availability: Parts Manufacturer Approvals1992 (Microfiche) AC 23-8A Flight Test Guide for Certification of Part 23 Airplanes c. Technical Standard Order (TSO) TSO-C26c Aircraft Wheels and Wheel-Brakes Assemblies, with Addendum I d. Industry Documents ARP 597C Wheels and Brakes, Supplementary Criteria Design for Endurance-Civil Transport Aircraft ARP 813A Maintainability Recommendations for Aircraft Wheels and Brakes AIR 1064B Brake Dynamics AS 1145A Aircraft Brake Temperature Monitor System ### Page 68 56 ARP 1619A Replacement and Modified Brakes and Wheels 2. BACKGROUND A review of recent replacement and modification wheel/brake system and installation approvals on part 23 airplanes has resulted in the need to provide FAA guidelines that clearly describe the changes and associated substantiation procedures involved. As contained herein, these guidelines will reflect upon issues that have been identified by industry under Aerospace Recommended Practice (ARP) 1619, revision A, which, in part, concerns the variance in compliance provisions associated with original equipment manufacturers (OEM) and non-OEM applicants. The guidelines will include a description of replacement or modified wheel, wheel/brake parts or assembly changes, a description and examples of associated "major" and "minor" changes, and a description of corresponding laboratory and airplane flight tests needed to ensure that requested changes will result in a continued level of airplane safety and performance. **3. CLASSIFICATION OF REPLACEMENT AND MODIFIED WHEEL/BRAKE** **CHANGES** a. Replacement of Wheel, Wheel/Brake Parts, or Assembly Changes A replacement wheel, wheel/brake part, or an assembly change is classified as one in which either the included parts or assemblies that are being changed are of equivalent design that will result in an equivalent level of certified type performance and safety to that exhibited by either the originally approved parts or assemblies. The change may be approved under a Parts Manufacturer Approval (PMA) by the provisions of part 21. Under part 21, § 21.303, an applicant may be eligible for approval of either PMA replacement parts or assemblies, or both, by demonstrating compliance in accordance with the following methods, as applicable: (1) Licensing Agreement Applicant should provide evidence of a licensing agreement or equivalent with the holder of a Technical Standard Order Authorization (TSOA), Type Certificate (TC), or a Supplemental Type Certificate (STC) together with the submission of any design data covered by the licensing agreement, as determined by the FAA. (2) Identicalness Applicant should provide evidence that the parts he produces will be identical in all respects to the corresponding parts of an approved 14 CFR type design, Technical Standard Order, or Parts Manufacture Approval. Data submitted should include all applicable design, material, and process specifications: that is, technical data that would specify all dimensions, tolerances, materials, processes, and specifications to the design of the corresponding part of an approved design. ### Page 69 Appendix 2AC 23-17A 57 (3) Airworthiness Requirements (Tests and Analyses) Applicant should provide evidence—in the form of drawings, test reports, computations, and other substantiating data—showing that the part meets either the applicable part 23 airworthiness requirements or the certification basis under which the airplane was approved. Compliance to applicable part 23 airworthiness requirements may include the following: (a) Part 23, § 23.55 and Accelerate-stop distance for Part 135, Appendix A commuter category airplanes and other airplanes that have accelerate-stop distance requirements, including airplanes that have published data such as stopping distances and brake energy/cooling charts in the Airplane Flight Manual. (b) Part 23, § 23.75 Landing. (c) Part 23, § 23.143 Controllability and maneuverability: General. (d) Part 23, § 23.231 Longitudinal stability and control. (e) Part 23, § 23.233 Directional stability and control. (f) Part 23, § 23.493 Braked roll conditions. (g) Part 23, § 23.731 Wheels. (h) Part 23, § 23.735 Brakes. (i) Part 23, § 23.1301 Function and installation. (j) Part 23, § 23.1309 Equipment, systems, and installations. (k) Part 23, § 23.1529 and Instructions for Continued Appendix G to Part 23 Airworthiness. (l) TSO-C26c Aircraft Wheels and Wheel-Brakes Assemblies, with Addendum 1. [Note: A description of the certification basis in which an airplane was approved can be obtained from the FAA.] ### Page 70 b. Identical Wheel, Wheel/Brake Parts, or Assembly An identical wheel, wheel/brake part, or assembly is classified as a replacement in which either the included parts or assemblies being changed are of an identical design and will result in an equivalent level of demonstrated performance to that exhibited by either the originally approved parts or assemblies. c. Modified Wheel, Wheel/Brake Parts, or Assemblies May Be Approved Under a Provision of Part 21 Under part 21, § 21.303, an applicant may be eligible for approval of modified wheel, wheel/brake parts, or assemblies by demonstrating compliance to methods identified under the above paragraph 3(a)(3) “Airworthiness Requirements (Tests and Analysis)”. d. Major and Minor Wheel, Wheel/Brake Parts, or Assembly Changes Since design changes appropriate to replacement and modified wheels, wheel/brake parts, or assemblies may involve changes to the original TSO wheel or wheel/brake assembly approval basis under an STC or TC in which the wheel or wheel/brake was installed, compliance to applicable provisions for "major" and "minor" design changes under part 21, § 21.93 or § 21.611, or both, should also be complied with per the following: (1) Major Design Changes A major design change to an existing TSO approved assembly is one that would require a substantially complete investigation of change for compliance to requirements under the TSO, and would result in a new type or model designation. A major design change to an airplane's type design or certification basis is one that could appreciably affect the weight, balance, structural strength, reliability, operational characteristics or other characteristics affecting the airworthiness of the airplane. Examples of such major design changes involving the wheel, wheel/brake parts, or assemblies include, but are not limited to, the following: (a) Structural material changes and friction material composition changes of heat sink elements that result in changes to FAA-approved performance data. (b) Reduction of original heat sink mass. (c) Change in the total brake actuation load or area. (d) Changes in the friction radius, the total number, or the area of friction faces or elements. (e) Fuse plug relocation in the wheels, change in release temperature, or a fuse plug redesign where a minor change has not been substantiated. ### Page 71 Appendix 2AC 23-17A 59 (f) Changes that would adversely affect the temperature-time profile of either a wheel or fuse plug, or both. (g) Relocation of "overpressure release" or "inflation valve." (h) Redesign of the wheel in a wheel/brake assembly, including a reduction in the wheel or brake structure that could adversely affect wheel strength or fatigue life. (i) Reduction in the wheel tie-bolt diameter or material strength of bolt and nut. (j) Change in the wheel bearing size that could or would adversely affect the wheel or bearing load capacity. (k) Changes in the wheel structural strength, deflection, fatigue life, or weight. (2) Minor Design Changes A minor design change to the TSO assembly (or airplane once the assembly is installed) is one that would have no appreciable affect on either the performance of the original TSO assembly or the certification basis (as identified above for major change) of the airplane in which the assembly is to be installed. Investigation into further compliance and FAA approval would normally be limited to minimal functional and compatibility tests. Original model numbers would be retained while part numbers could be used to identify minor changes for TSOA. See paragraph 6.b. “Part Numbering.” of this AC in § 23.735 Brakes, for PMA part numbering requirements. Examples of potential minor design changes involving wheel, wheel/brake parts or assemblies might include but are not limited to the following: (a) Brake friction material changes or heavier heat sink elements that do not result in a change of FAA-approved performance data. (b) Structural improvements to improve fatigue life. (c) Paint/corrosion protection changes. (d) Changes to bleed ports or tube and service fittings. (e) Revised over-inflation devices. 4. SUBSTANTIATION PROCEDURES Replacement and modified wheels, wheel/brake parts or assembly changes should be substantiated by conducting the necessary analytical investigations, laboratory testing, or airplane testing, or all of these, to ensure that the change can be made without adversely affecting aircraft safety and associated braking and rolling performance. A substantiation plan should first be proposed by the applicant for FAA approval followed by the applicant's implementation of the plan. ### Page 72 a. Substantiation Plan A proposed substantiation plan may be presented to the FAA for approval that identifies the applicant's requested change and intended approach in substantiating the change in accordance with the methods addressed under this section. The plan should include the following: (1) A description of the replacement or modified part or assembly, or both. (2) An assessment covering the applicable airworthiness requirements involved. (3) A statement of change that is determined to be either "major" or "minor" along with the basis for the classification relative to the applicable requirements of part 21. (4) An assembly drawing reflecting the replacement or modified part or assembly, or both. (5) Aircraft installation drawings/instructions. (6) The substantiation method, which includes an analysis/test protocol. (7) The method of identification and maintenance procedures that will be utilized. (8) The quality management and quality assurance system under which either the part or assembly, or both, will be produced. b. Substantiation Requirements As contained herein, the recommended substantiation requirements for replacement and modified wheels, wheel/brake parts or assemblies are based upon changes for which approval is requested, and the impact a new part or assembly will have on prior certification. If the replacement and modified wheel, wheel/brake part or assembly meets the minimum applicable airworthiness requirements to the product (airplane) on which either the part or assembly, or both, is to be installed, but not the Airplane Flight Manual (AFM) performance data, then the applicant should provide the applicable performance data in an FAA approved AFM or AFM supplement. Depending upon the type and extent of change (as defined under “Section 3, CLASSIFICATION OF REPLACEMENT AND MODIFIED WHEEL/BRAKE CHANGES”) and either engineering or pilot judgment, or both, FAA approval will be determined on the basis of compliance with the following substantiation requirements: ### Page 73 Appendix 2AC 23-17A 61 (1) Replacement Wheel, Wheel/Brake Parts or Assemblies (a) Brake-Anti-skid Compatibility Replacement part or assembly changes, or both, defined under paragraph 3(d)(2) titled “Minor Design Changes,” are considered to be minor whether they are proposed by the original wheel and brake manufacturer who holds the TSO authorization or by another manufacturer seeking to produce a replacement part or assembly. While such changes are not expected to affect braking performance, functional landings may be required as a minimum to verify airplane/pilot/brake/anti-skid combination compatibility (reference part 23, § 23.735(d)). Normally five (5) non-instrument, functional landings are necessary to verify this compatibility. (b) Brake Rotors/Stators In general, changes to the friction surfaces of the aircraft brake, including the stator and rotor, are considered to represent a major change per 3(d)(1), titled “Major Design Changes,” unless it can be shown that the change cannot affect the airplane stopping performance, brake energy absorption characteristics, or continued airworthiness (reference part 23, § 23.735(a)/(e)). In addition, if changes in heat sink friction components are proposed, certain provisions of 4(b)(2), titled “Modified Wheel, Wheel/Brake Parts, or Assemblies,” may also be applicable. Changes to continuing airworthiness, such as thermal control, vibration control, etc., should also be considered for the major/minor determination. In this regard, the original manufacturer of the wheel or wheel/brake assembly who holds the TSO authorization may possess data sufficient to show that such changes could be considered minor (e.g., airplane performance would not be affected). On the other hand, a manufacturer other than the original manufacturer who may wish to produce replacement rotors and stators may not have data sufficient to show that performance would not be affected. In this case, the major/minor status would be determined by applicable dynamometer tests per TSO-C26c and some functional airplane tests as a minimum. (c) Brake Performance Equivalency It may be difficult to determine identicalness but a finding of equivalency can be shown by additional design, analysis, and dynamometer tests as applicable. A change to an approved part that is determined to be minor can be validated on the dynamometer by a controlled test at the maximum certified kinetic energy capacity of the original brake assemblies from TSOA, or the dynamometer testing may be done to the design landing and accelerate-stop kinetic energy levels appropriate to the aircraft [Reference part 23, § 23.735 (a) through (e)]. The following dynamometer test protocol is acceptable to validate replacement rotors/stators proposed by an applicant other than the original TSO holder: ### Page 74 1. Use of new stator or rotor parts in the replacement manufacturer's brakes for each dynamometer test will be required in order to minimize test configuration variables. If rebuilt or in-service components other than these fail during testing, it should be realized that the results may be questionable. Suspect tests would be carefully scrutinized by the FAA, and retesting may then be necessary. Test methods, test hardware (including the tire size, ply and condition), and test procedures should be the same to ensure proper comparative evaluations. If brake friction materials are being compared, the heat sinks to be used for maximum certified kinetic energy (KE) testing should not have been subjected to test energies higher than design landing energy. 2. The maximum certified kinetic energies approved under TSOA for the original manufacturer are proprietary data. Therefore, a PMA applicant that desires to maintain the TSOA status of a modified assembly will have to do the testing in this paragraph without knowing the kinetic energy levels the OEM tested for the TSOA. A series of tests may be necessary for a replacement manufacturer to reach the maximum certified level of the original manufacturer's brake. For each succeeding run, the KE will be increased by at least 5 percent over the previous run until the maximum certified KE level is reached. The initial KE level for this series of tests will be at the discretion of the applicant. If maintaining the TSOA is not desired, the PMA applicant may perform dynamometer tests at the airplane derived kinetic energy levels. 3. Maximum braking force pressure, derived from the airplane maximum brake pressure capability, is to be applied during the tests. 4. Fuse plugs may be released or the tire deflated after each test run to reduce the risk to test personnel. 5. A minimum of five (5) functional landings for anti-skid equipped airplanes and a minimum of three (3) functional landings for non-anti-skid equipped airplanes, as described above, are needed. (d) Worn Brakes [optional] While there are no provisions under part 23 to require the evaluation of brake performance using worn brakes, there have been rejected takeoff accidents in which the brakes on subject airplanes were at or very near their completely worn state of energy absorption capability and stopping capability. Therefore, as an optional test to a replacement brake performance evaluation (when there is question concerning variances in worn brake performance), it is recommended that such an assessment on the dynamometer be undertaken to support compliance with maximum Rejected Takeoff (RTO) performance in the AFM. Dynamometer tests simulating a maximum energy RTO should be performed on the replacement brake assemblies with individual brakes within 10 percent of their wear limit (e.g., at least 90 percent worn). The tests, used to verify the safety of a replacement brake system and to determine the maximum energy ### Page 75 Appendix 2AC 23-17A 63 absorption capability of brakes in their fully worn state, should be substantiated as being representative of actual airplane and runway conditions. (2) Modified Wheel, Wheel/Brake Parts, or Assemblies (a) Modified Brake Design This laboratory and airplane test requirement applies to the addition of a major change brake design to an existing airplane for which FAA approved braking performance test data exists. Testing may be performed either for performance credit or to the existing performance level of the aircraft. As provided under examples of 3(d)(1), a modified brake is one that contains new or modified parts that may cause a significant variance in the kinetic energy absorption characteristics, AFM stopping distances and continuing airworthiness of the brake [Reference 14 CFR § 23.735 (a) through (e)]. Substantiating laboratory and airplane flight testing required for approval of a major changed brake will include the following: 1. For improved performance (aa) Applicable dynamometer tests under TSO-C26c. (bb) Instrument flight tests to include six (6) takeoffs and six (6) landings. The six landings are to be conducted on the same wheels, tires, and brakes. All tests should be conducted with engines trimmed to the high side of the normal idle range, if applicable. The engine idle schedule may include a flight-idle schedule which may be applicable to the test. (cc) Additional tests may be necessary for each airplane configuration change (e.g., takeoff flaps, landing flaps, nose wheel brakes, anti-skid devices inoperative, deactivation of wheel/brakes, etc.). (dd) Brake system response evaluation. (ee) Parking brake adequacy. Tires are allowed to skid during maximum power engine checks. (ff) Alternate braking system stops. (gg) Fuse plug evaluation. (hh) Anti-skid compatibility on wet runway. (ii) Taxi tests, to ensure that ground handling, maneuvering, and brake sensitivity are satisfactory, should be conducted. ### Page 76 64 (jj) At least two (2) braking stops, one at maximum takeoff weight and one at minimum landing weight, should be conducted on a wet runway to verify brake and anti-skid compatibility. [Note: Improved performance implies an increase in the friction coefficient (mu) versus energy level for the desired operation(s) and may be requested for landing, RTO's, or a specific configuration such as anti-skid "on" only.] 2. For equivalent performance (aa) Applicable dynamometer tests under TSO-C26c. (bb) A sufficient number of conditions to verify the existing approved performance levels (RTO and landing for either TSOA levels or Airplane Flight Manual levels). Consideration should be given to verification of fuse plugs, performance verification at appropriate energy levels, and configuration differences, including anti-skid on and off. (cc) Taxi tests, to ensure that ground handling, maneuvering, and brake sensitivity are satisfactory, should be conducted. (dd) At least two (2) braking stops, one at maximum takeoff weight and one at minimum landing weight, should be conducted on a wet runway to verify brake and anti-skid compatibility. [Note: Equivalent performance implies that sufficient data will be obtained to verify that the performance level for the change is equal to or better than the existing performance levels. The change may be for the purpose of changing the c.g. envelope, or for airplane configuration changes (such as flap angles), and may apply to specific operations (such as landings).] 3. For extended performance (aa) Applicable dynamometer tests under TSO-C26c. Consideration should be given to the items in Section 4b(2)(a)(2). (bb) A sufficient number of conditions to define the extended life and determine equivalency to the existing performance levels. Consideration should be given to the items in Section 4b(2)(a)(2). [Note: Extended performance implies that the existing certification µ versus energy line will extend to establish the braking force level for a proposed change, such as gross weight or the desired maximum energy level, and may be applied to a specific operation (such as landing only).] ### Page 77 Appendix 2AC 23-17A 65 (b) Modified Anti-skid System This airplane test requirement applies to the addition of a new anti-skid system or changes to an existing anti-skid system that may affect airplane performance (e.g., new anti-skid system, or a change from couple to individual wheel control). A sufficient number of either airplane performance tests or functional tests, or both, should be conducted to verify existing approved performance anti-skid "on" levels. In the event an increase of braking performance is desired, full airplane performance testing will be required [reference part 23, § 23.735 (d)]. (c) Modified Fuse Plugs/Wheels This item covers the addition of a significant modification to any portion of the existing wheel design on an airplane (change of wheel design, redesign, or relocation of fuse plugs). The following airplane tests can be performed when such changes are made: 1. One airplane braking test should be conducted to show that the fuse plugs will release when excessive energies are absorbed. 2. One airplane braking test should be conducted to verify the maximum kinetic energy at which fuse plugs will not release (fuse plug substantiation). Dynamometer tests are not adequate for this test. [Note: Wheel fuse plug integrity should be substantiated during braking tests where the energy level simulates the maximum landing energy. It should be demonstrated that the wheel fuse plugs will remain intact and that unwanted releases do not occur. One acceptable method to determine this is as follows:] (aa) Set engine idle thrust at the maximum value specified (if applicable). The engine idle schedule may include a flight-idle schedule which may be applicable to the test. (bb) Set tire pressures to the minimum value appropriate for the airplane test weight. (cc) Taxi at least three miles (normal braking, at least three intermediate stops, and all engines operating). (dd) Conduct accelerate-stop test at maximum landing energy, maintaining the deceleration rate consistent with the values used to determine performance distance. (ee) Taxi at least three miles (normal braking, at least three intermediate stops, and all engines operating). ### Page 78 66 (ff) Park in an area to minimize wind effects until it is ensured that fuse plug temperatures have peaked and that no plugs have released. Instead of simulating the maximum kinetic energy landing during an accelerate-stop test, an actual landing and quick turnaround may be performed; however, caution should be exercised in order to prevent jeopardizing the safety of the flight crew and airplane if the wheel plugs release right after liftoff, requiring a landing to be made with some flat tires. The following elements should be included in the tests: (aa) Set engine idle thrust at maximum value specified (if applicable). (bb) Set tire pressures to the minimum value appropriate to the airplane test weight. (cc) Conduct a landing stop at maximum landing energy, maintaining the acceleration rate consistent with the values used to determine performance distance. (dd) Taxi to ramp (three miles minimum with normal braking, at least three intermediate stops, and all engines operating). (ee) Stop at the ramp. Proceed immediately to taxi for takeoff. (ff) Taxi for takeoff (three miles minimum with normal braking, at least three intermediate stops, and all engines operating). (gg) Park in an area to minimize wind effects until it is ensured that fuse plug temperatures have peaked and that no plugs have released. Fuse plug protection of wheels and tires should be demonstrated to show that the fuse plugs will release when excessive energies are absorbed. Normally, this will occur during RTO performance tests. (d) Accelerate Stop Tests Accelerate-stop tests for commuter category airplanes and other airplanes are defined under part 23, §§ 23.55 and 23.735 (e). Accelerate-stop tests should be conducted for all modified wheel, wheel/brake parts or assemblies involving a major design change when this testing was performed for the certification of the original brake assembly. Such tests should include substantiation of the critical maximum brake energy stop (highest ground speed based on the V1 speed applicable to the maximum altitude and temperature the airplane is certified for according to the FAA approved AFM). On airplanes with wheel fuse plugs, a satisfactory demonstration of fuse plug compatibility should be conducted as stated under Item (2)(c) titled, “Modified Fuse Plugs/Wheels.” ### Page 79 Appendix 2AC 23-17A 67 (e) Other Substantial Airplane Tests Depending upon the extent of wheel, wheel/brake part, or assembly modifications that may be involved, there will be a number of airplane tests that should be considered in addition to those above. As applicable to specific changes and to the type of part 23 airplane involved, the following are tests that may be appropriate and required by the FAA for approval: 1. Brake KE Absorption Tests Verify that the brake kinetic energy absorption test determined by the laboratory test meets the TSO requirements and the airplane manufacturer's requirements (to be identified by the FAA) [reference part 23, § 23.735(a) through (e)]. The engine idle schedule may include a flight-idle schedule which may be applicable to the test. 2. Brake Pressure Test Verify brake pressure tests conducted under the TSO are adequate for the brake system pressure on the airplane, as determined by the manufacturer's brake system pressure data. Conduct a brake pressure test on the airplane if manufacturer's brake system data is not available to verify the adequacy of the TSO test [reference part 23, § 23.735(c)]. 3. Taxi Ground Handling Tests Perform taxi tests to ensure that ground handling, controllability, maneuverability, and brake sensitivity are satisfactory. Use normal braking, intermediate stops, with all engines operative [see 4b(2)(a), Modified Brake Design]. 4. Wet Running Tests Perform brake stops on a wet runway to verify brake and, if applicable, anti-skid system compatibility [see 4b(2)(a), Modified Brake Design]. 5. Function Reliability Tests Perform function reliability landing stops. Normally six maximum brake landings should be satisfactorily conducted on the same set of wheels, tires, and brakes [see 4b(2)(a), Modified Brake Design]. 6. Landing Performance Tests Determine that the landing performance is adequate to the previously approved performance data shown in the AFM. If the AFM performance data is not available because that it is not required by the airplane certification basis, the manufacturer's data ### Page 80 68 (if available) provided to the pilot should be used as a basis for comparison [see 4b (2)(a) Modified Brake Design]. 7. Static Torque Tests Determine whether there is adequate static torque when parked and during appropriate engine run up conditions [reference part 23, § 23.735(b)]. 8. Brake Response Tests During the aforementioned tests, brake response characteristics should be monitored for unacceptable vibrations, squeal, fade, grabbing, and chatter. These characteristics may have a destructive effect on the brake assembly components and may be pertinent to endurance of landing gear system components. 5. INSTRUCTIONS FOR CONTINUED AIRWORTHINESS A PMA applicant may be required to furnish instructions for continued airworthiness if the article on which the part is eligible for installation has an existing set of instructions for continued airworthiness that are not considered adequate for the applicant's PMA part (reference part 23, § 23.1529). **6. IDENTIFICATION OF PMA PARTS** a. General Under part 45, § 45.15, parts produced under a PMA should be permanently and legibly marked in a manner that will enable persons to identify the following: • It is a PMA part. • The manufacturer. • The part number. • The type certificated product(s) or TSOA article(s) on which it may be installed. For a part based on an STC, the identification of installation eligible type certified products should include reference to the STC. In accordance with part 45, § 45.14, parts that have been identified as critical components should be marked with a part number, or equivalent, and serial number or equivalent. If the TC or TSOA holder applies serial numbers to a critical part, the PMA holder should also "permanently mark" their parts with serial numbers. Note: Due to the harsh environments that wheels and brakes experience, decals or adhesive backed "metalcals" are not considered permanent forms of marking. Metal stamping, etching or permanently affixing a data plate with rivets or drive screws in a non-critical area is satisfactory. Laser marking is also acceptable if it can be read under 2X magnification. Ink stamping is allowed only if more permanent means are not possible. ### Page 81 Appendix 2AC 23-17A 69 b. Part Numbering The PMA holder's part should be numbered such that it is sufficiently different from the OEM holder's part number to be distinguishable. The OEM holder's part number with a prefix/suffix is sufficient for this purpose. The requirements of part 45, § 45.15(a)(2), to mark with name, trademark, or symbol of the PMA holder may be satisfied by the prefix/suffix if the prefix/suffix is done consistently across the PMA holder's product line. The FAA-PMA letter should show the type approved part number with which the PMA holder's part is interchangeable. c. Parts Manufactured Under License When the PMA is issued by showing evidence of a license agreement or equivalent, the PMA part number may be identical to that on the type certificated part providing the PMA holder also meets the requirements of part 45, § 45.15(a)(1) and (2) to permanently mark the part with the letters "FAA-PMA" and the name, trademark, or symbol of the PMA holder. d. Parts that are Impractical to Mark In all cases where the part is found by the FAA to be too small (or to have other characteristics that make it impractical) to mark all (or any) of the information on the part, the information not marked on the part should be put on the tag that is attached to the part or marked on the container for the part. If the number of certificated products or TSOA articles on which the part is eligible for installation is too long to be practicable to include with the part, the tag or container may refer to a readily available manual or catalog made available by the PMA holder for part eligibility information. Amendment 23-34 and Subsequent Policy for 14 CFR part 23, § 23.735, Brakes, and Specific Sections versus § 23.1309; Equipment, Systems and Installations Brake systems are approved based on compliance to the specific standard, § 23.735, not the general standard, § 23.1309. This is true for any installations where there are specific standards in 14 CFR part 23 (i.e., brakes, autopilots, ice protection systems, etc.). In the past, the Small Airplane Directorate has not considered brake failures catastrophic or severe-major due to mitigating factors, which include the following: low stall speeds, minimum field lengths for landing, and propeller feathering or reverse thrust. For commuter category airplanes where loss of the brakes is catastrophic or severe-major, we will accept an emergency brake system or a single failure in the primary system that meets the following requirements (based on § 25.735(b)): ### Page 82 70 In the following circumstances, the brake system and associated systems must be designed and constructed so that the airplane may be brought to rest under conditions specified in § 23.75: - If any electrical, pneumatic, hydraulic, or mechanical connecting or transmitting element (excluding the operating pedal or handle) fails, or - If any single source of hydraulic or other brake operating energy supply is lost. The airplane can be brought to rest with a mean deceleration during the landing roll of at least 50 percent of that obtained in determining the landing distance as prescribed in that section. Subcomponents in a brake assembly such as brake drums, shoes and actuators (or their equivalents), shall be considered as connecting or transmitting elements unless it is shown that leakage of hydraulic fluid from a failure of sealing elements in these subcomponents would not reduce the braking effectiveness below that specified in this paragraph. We agree that requiring an emergency brake system is not desirable for small single-engine airplanes. We cannot agree that brake system failure for commuter category and turbojet powered airplanes can automatically be classified as severe-major. Mitigating factors must always be considered for any part 23 airplane. JAA ACJ 23.735(c) is acceptable for FAA certification. ### Page 83 23.737 Skis Amendment 23-45 and Subsequent See § 23.505, Supplementary conditions for ski-planes, for additional guidance about aircraft skis. ### Page 84 23.745 Nose/tail wheel steering No policy available as of June 30, 2001. ### Page 85 **FLOATS AND HULLS** 23.751 Main float buoyancy Original Issue and Subsequent For a twin float seaplane or amphibian aircraft, the 80 percent excess buoyancy requirement should be applied to both rather than each float. The rules for twin float aircraft do not address water stability or capsizing. They only require that the aircraft remain afloat with any two compartments of the main floats flooded. The history does not support the position that the aircraft will remain afloat indefinitely without capsizing with two compartments flooded. However, if an unsafe condition exists, certification should be denied under the provisions of part 21, § 21.21(b)(2). An example of an unsafe condition could be capsizing so rapidly that the occupants could not safely exit. Capsizing that is delayed long enough to permit taxi to the shore or dock is not an unsafe condition. The time to capsize should be listed in the Emergency Procedures Section of the Airplane Flight Manual. ### Page 86 23.753 Main float design No policy available as of June 30, 2001. ### Page 87 23.755 Hulls Original Issue The discussion on capsizing and unsafe conditions in 23.751 applies to this rule also. Amendment 23-45 and Subsequent This amendment changed the rule to prohibit capsizing in fresh water for planes over 5,000 pounds with two adjacent compartments flooded, and those between 1,500 and 5,000 pounds with any single compartment flooded. ### Page 88 23.757 Auxiliary floats No policy available as of June 30, 2001. ### Page 89 **PERSONNEL AND CARGO ACCOMMODATIONS** 23.771 Pilot compartment No policy available as of June 30, 2001. ### Page 90 23.773 Pilot compartment view Amendment 23-14 and Subsequent See AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes, for guidance on this regulation. ### Page 91 23.775 Windshields and windows Original Issue and Subsequent The rule requires the luminous transmittance (LT) be no less than 70 percent when the pilot is seated in a normal flight position. This rule does not specify how the LT is to be measured. Industry, federal practices, standards and airframe manufacturer specifications have specified a minimum LT measurement per Federal Standard 406, Method 3022 or equivalent. On the basis of available data, we cannot determine that an LT of 70 percent is in itself an unsafe condition. There are other factors such as windshield/window inclination from vertical. The criteria to determine an unsafe condition is a qualitative pilot evaluation. Nonsplintering material in this section refers to materials such as safety glass, which shatter into small fragments that can cause superficial injuries. Splintering materials are those that shatter into slivers, which can result in serious injuries. Policy for 14 CFR part 23, § 23.775, Temperature and Temperature Gradient Requirement Certification: Section 23.775 requires the following: - “The design of windshields, windows, and canopies in pressurized airplanes must be based on factors peculiar to high altitude operation, including – The effects of temperatures and temperature gradients;” and - “If certification for operation above 25,000 feet is requested, the windshield, window panels, and canopies must be strong enough to withstand the maximum cabin differential loads combined with critical aerodynamic pressure and temperature effects, after failure of any load-carrying element of the windshield, window panel, or canopy.” On-airplane testing is one acceptable Means of Compliance (MOC) to these requirements. Other acceptable MOC’s are component level testing, a combination of tests and analyses, or an analysis based on known temperature and temperature gradient effects for a specific material and installation. Qualification by similarity is also acceptable. Note: At this time, we are not aware of any case where an applicant has used an analysis without testing, but we should consider this if it is proposed by an applicant. ### Page 92 80 The testing or analysis would have to consider pressurization loads or fail-safe loads with the temperature and temperature gradient effects. These would be in addition to the cyclic or fail-safe testing of the complete pressure cabin at ambient temperature. Current practice is to superimpose the pressure loading and the thermal gradients in a test. The testing and conditions vary from full-scale test articles with environmental chambers around the windows and windshields to component tests. For fail safe requirements, most manufacturers are using a “failed” article (cut before test) while running a test to fail-safe loads or better. If the bird strike requirements in § 23.775(h)(1) are applicable (Amendment 23-49), the applicant should consider the effects of temperature and temperature gradients when showing compliance to the two-pound bird strike. JAA ACJ’s 23.775(f) and 23.775(g), which apply to § 23.775(g) and (h) in 14 CFR part 23, are acceptable for FAA certification. ### Page 93 23.777 Cockpit controls Original Issue and Subsequent The FAA has no rule preventing placement of non-flight controls on control wheels, but we consider such installations to be marginal since a switch could easily be confused with microphone or autopilot switches. Approval would require special crew training and AFM guidance to ensure it was used properly. We strongly recommend that all redundant cockpit controls be symmetrical from one side of the cockpit to the other. Amendment 23-33 and Subsequent Section 23.777(c)(4) was added by this amendment. It requires that airplanes with sideby-side pilot seats and two sets of powerplant controls, have one set on the left console and one set on the right console. We will consider an Equivalent Level of Safety Finding for airplanes with one set on the left hand side and one on or near the cockpit center line. The preamble to the 23-33 Amendment shows that floor mounted, mechanical flap controls are acceptable. 14 CFR part 23/CAR 3 Airplanes; Clarification of Type Certification Process of Single Lever Power Controls The purpose of this guidance is to provide recommendations for certification of single lever power controls (SLPC) installed in part 23/CAR 3 airplanes. There are airplanes certificated or currently undergoing certification that have combined the features of two or more of the cockpit powerplant controls for power (thrust), propeller (rpm control), and mixture control (condition lever and fuel cutoff for turbine powered airplanes) into a single power lever. The design feature of an SLPC was not envisioned by part 23/CAR 3. Further, an SLPC cannot meet the standards imposed by §§ 23.777(d) and 23.781(b) as amended by Amendment 23-33. The current amendment level of part 23 (Amendment 23-53) contains regulations that allow evaluation of an SLPC without the need for Special Conditions (e.g., §§ 23.777(a)(b), 23.779(b)(1), 23.1309). However, since an SLPC was not envisioned at the time Amendment 23-33 was adopted, the question of compliance with §§ 23.777(d) and 23.781(b) as amended by Amendment 23-33 still exists. ### Page 94 82 Due to recommendations made by the NTSB, Amendment 23-33 provided specific location, height, and shape requirements for a number of cockpit controls, including power, propeller, and mixture controls. With the design feature of an SLPC integrating the functions of multiple controls into a single cockpit control, a nonstandardized design approach for the affected powerplant cockpit controls is used. Additionally, § 23.1141(a) states: “Powerplant controls must be located and arranged under § 23.777. . . .” However, an SLPC, as described earlier, cannot be arranged in accordance with § 23.777; therefore, compliance with § 23.1141(a) is not possible. Notice No. 84-12, which was the basis for Amendment 23-33, described the intent of §§ 23.777 (d) and 23.781 (b). As stated in the notice: "An effective means of enhancing pilot experience and training would be to require complete standardization in cockpit design. While such action may initially improve the level of safety, it might ultimately inhibit design advancement and result in lower levels of safety than would have evolved without such a total standardization. An effective and practical means of enhancing the effectiveness of pilot training and enhancing safety would be to require standardization of location, shape, color, and direction of movement of those cockpit controls. This would have minimal adverse effect on design advancement." From the preceding, it is obvious that the FAA and industry did not envision or address the future use of an SLPC when drafting this rulemaking, but it was intended to allow design advancements that would enhance safety. An SLPC is a design advancement in the public interest and does not adversely affect safety. Therefore, an SLPC will meet the intent, but not literal compliance, of §§ 23.777(d) and 23.781(b) as amended by Amendment 23-33. We, therefore, recommend use of an Equivalent Level of Safety finding for airplanes with a certification basis of Amendment 23-33 or later when making compliance determinations for §§ 23.777(d) and 23.781(b). For these airplanes, Special Conditions are usually unnecessary. In some cases, however, the applicable airworthiness standards may not be adequate due to other novel or unusual features of the aircraft, and Special Conditions may be warranted. For airplanes with a certification basis prior to Amendment 23-33, no special considerations will be needed unless they involve other novel or unusual design features not covered by the applicable regulations. ### Page 95 83 We have initiated regulatory action to revise part 23 to allow incorporation of an SLPC without special considerations. However, until these actions have been completed, the recommendations in this advisory circular may be used for certification of an SLPC on part 23/CAR 3 airplanes. ### Page 96 23.779 Motion and effect of cockpit controls Original Issue—Reserved The original rule only specified primary aerodynamic and throttle motion. Amendment 23-33 and Subsequent This amendment requires that the propeller control should move forward to increase rpm. Therefore, a propeller pitch control on the vertical instrument panel does not comply. A switch located in a horizontal position with forward motion to increase rpm does comply. Other designs would have to be considered by an Equivalent Level of Safety Finding. We would consider a propeller pitch control switch that would increase propeller rpm when moved to the “UP” position with a placard to denote rpm change to be eligible as an equivalent. We have no objection to the propeller pitch control switch being spring loaded against the fine and coarse propeller blade angles. The airplane should be evaluated to ensure no unsafe operating condition occurs with a propeller switch in each critical blade position. See § 23.777 for guidance applicable to single power lever controls. ### Page 97 23.781 Cockpit control knob shape Amendment 23-33 and Subsequent This rule did not envision a single power lever installation. See § 23.777 for guidance applicable to single power lever controls. ### Page 98 23.783 Doors Original Issue and Subsequent An Equivalent Safety Finding process for part 23, § 23.783(b), should include the following: a. Pilot operated locks when the propeller stops turning. b. A special operating procedure to ensure the door is opened only after the propeller has stopped turning should be provided in the flight manual and on the inside of the door. c. If an electric lock is used, complete loss of electric power should not affect opening the door. d. The door should be designed and placarded so it can be opened from the inside by passengers and from outside by ground personnel. e. A railing or guard that would deploy to guide passengers away from the propeller plane should be provided as an integral part of the door. f. If engagement of the engine starter would be an immediate hazard to a person near the propeller, an interconnection between the door and the engine starting circuit should be included in the design. Amendment 23-34 and Subsequent The direct visual inspection of the locking mechanism by crewmembers to determine whether external doors, for which the initial opening movement is outward, are fully closed and locked may be conducted from outside the airplane. It will be necessary to provide a means to visually inspect each individual lock of the locking mechanism. Means that do not permit direct visual inspection of each lock are unacceptable unless there is no failure mode of the locking mechanism that would allow a false visual indication that each latch is properly positioned and locked. If optical devices are used from either inside or outside, it should be determined that they are not subject to fogging, to obstruction by foreign objects, or to a false indication of a locked condition. The locking mechanism should incorporate features that provide a positive means to prevent the door from vibrating open throughout the approved operating envelope. Over center features of the mechanism are not acceptable as a locking means. Also, it should not be possible to position the locks in a locked position if any of the latches are not in the fully latched position. ### Page 99 87 Amendment 23-36 and Subsequent Section 23.783(c) was adopted to provide standards that would ensure the opening means of passenger and crew doors were simple, easy to locate, could be operated in darkness, and ensure the doors met particular marking requirements. Overly complex opening means had been identified as a major contributor in accident investigations. The particular marking means are those of § 23.811. Paragraph 23.783(c)(3) was adopted mainly as a measure to ensure that the opening means of passenger and crew doors were kept as simple as possible, and that these doors could be located and opened in a timely manner. Paragraph 23.783(c)(4) was adopted in order to make the location of cabin doors more conspicuous and to facilitate emergency evacuation. These requirements do not mandate the use of self-illuminated or electrically illuminated external markings. A reasonable and acceptable method of compliance can be found in § 23.807 by substituting “passenger or crew door” where reference is made to “emergency exit.” This amendment added a requirement that doors be reasonably free of jamming from fuselage deformation from the ultimate load factors found in the General Requirements for Emergency Landing Conditions (23.561) and the limit Proof of Structure (23.307(a)). Some applicants choose ultimate Proof of Structure loads as a higher level of safety than proposed here. Methods used to demonstrate compliance with these requirements should include: A determination from tests that each emergency door can be opened from inside and outside the airplane after imposing each critical static test load condition from the following limit flight and ground design load requirements (provided they equal or exceed the Emergency landing Conditions). FLIGHT LOADS General (§ 23.321) Symmetrical flight conditions (§ 23.331) Flight envelope (§ 23.333) Design airspeeds (§ 23.335) Limit maneuvering load factors (§ 23.337) Gust load factors (§ 23.341) Pressurized cabin loads (for flight, § 23.365(a) and (b); and for landing, § 23.365(a) and (c)) ### Page 100 GROUND LOADS General (§ 23.471) Ground load conditions and assumptions (§ 23.473) Landing gear arrangement (§ 23.477) Level landing conditions (§ 23.479) Tail down landing conditions (§ 23.481) One-wheel landing conditions (§ 23.483) Side load conditions (§ 23.485). Amendment 23-49 and Subsequent This amendment adds a requirement that passenger doors not be located with respect to any other potential hazard as well as the propeller disk. These hazards could include hot anti-ice, hot de-ice surfaces, and sharp objects on the airplane structure. ### Page 101 23.785 Seats, berths, litters, safety belts, and shoulder harnesses Original Issue and Subsequent Methods of Approval of Retrofit Shoulder Harness Installations in Small Airplanes A retrofit shoulder harness installation in a small airplane may receive approval by Supplemental Type Certificate (STC), Field Approval, or as a minor change. An STC is the most rigorous means of approval and offers the highest assurance the installation meets all the airworthiness regulations. A Field Approval is a suitable method of approval for a shoulder harness installation that needs little or no engineering. Shoulder harness installations may receive approval as a minor change in certain cases. In such cases, the FAA certificated mechanic who installs the shoulder harness records it as a minor change by making an entry in the maintenance log of the airplane. The FAA does not encourage the approval of retrofit shoulder harness installations as minor changes. The preferred methods of approval are Supplemental Type Certificate or Field Approval. However, the FAA should not forbid the approval of a retrofit shoulder harness installation as a minor change in: the front seats of those small airplanes manufactured before July 19, 1978, and in other seats of those small airplanes manufactured before December 13, 1986. A retrofit shoulder harness installation may receive approval as a minor change in these small airplanes if: The installation requires no change of the structure (such as welding or drilling holes). The certification basis of the airplane is 14 CFR part 23 before Amendment 23-20, part 3 of the Civil Air Regulations, or a predecessor regulation. In addition, a minor change installation should follow the guidance for hardware, restraint angles, and attachment locations provided in the following: Advisory Circular (AC) 43.13-2A, Acceptable Methods, Techniques, and Practices Aircraft Alterations. AC 21-34, Shoulder Harness - Safety Installations. AC 23-4, Static Strength Substantiation of Attachment Points for Occupant Restraint System Installations. ### Page 102 90 Installations approved as a minor change may not provide the occupant with the protection required by regulation (Civil Air Regulation (CAR) 3.386 or 14 CFR part 23, § 23.561). However, a properly installed retrofit shoulder harness installation is a safety improvement over occupant restraint by seat belt alone. References Advisory Circular (AC) 21-34, Shoulder Harness — Safety Belt Installations, June 4, 1993. AC 23-4, Static Strength Substantiation of Attachment Points for Occupant Restraint System Installations, June 20, 1986. AC 43.13-2A, Acceptable Methods, Techniques, and Practices — Aircraft Alterations, Revised 1977. Order 8300.10, Airworthiness Inspector’s Handbook, Change 12, December 14, 1999, Volume II. Technical Standard Order (TSO)-C114, Torso Restraint Systems, March 27, 1987. Technical Standard Order C-22g, Safety Belts, March 5, 1993. What are the Requirements? 1. Front seat shoulder harnesses required. Section 23.785 of 14 CFR part 23 as amended by Amendment 23-19 effective July 18, 1977, required all normal, utility, and acrobatic category airplanes for which application for type certificate was made on or after July 18, 1977, to have an approved shoulder harness for each front seat. Section 91.205(b)(14) requires all small civil airplanes manufactured after July 18, 1978, to have an approved shoulder harness for each front seat. The shoulder harness must be designed to protect the occupant from serious head injury when the occupant experiences the ultimate inertia forces specified in § 23.561(b)(2). The inertia force requirements are discussed in paragraph 3 below. 2. Shoulder harnesses required at all seats. Section 91.205(b)(16) requires all normal, utility, and acrobatic category airplanes with a seating configuration of 9 or fewer seats, excluding pilot seats, manufactured after December 12, 1986, to have a shoulder harness, for forward-facing and aft-facing seats, that meets the requirements of § 23.785(g) [which requires that the occupant be protected from the ultimate inertia forces specified in § 23.561(b)(2)]. Section 23.785(g) also ### Page 103 91 provides: “For other seat orientations, the seat and restraint means must be designed to provide a level of occupant protection equivalent to that provided for forward and aft-facing seats with safety belts and shoulder harnesses installed.” The above part 91 operating rule stems from § 23.2, Special retroactive requirements, Amendment 23-32, effective December 12, 1985. 3. Belts or harnesses provided for in the design. Civil Air Regulation (CAR) 3.386 and part 23, § 23.561, Amendments 23-0 through 23-34, effective February 17, 1987, require occupant protection from serious injury during a minor crash landing when “proper use is made of belts or harnesses provided for in the design,” when the occupants are subjected to the following ultimate inertia forces: Normal & Utility Category Acrobatic Category Forward 9.0g 9.0g Sideward 1.5g 1.5g Upward 3.0g 4.5g With Amendment 23-36, effective September 14, 1988, the text of § 23.561 quoted above was changed to read: “proper use is made of seats, safety belts, and shoulder harnesses provided for in the design.” Section 23.785(b) was also changed to read: “Each forward-facing or aft-facing seat/restraint system in normal, utility, or acrobatic category airplanes must consist of a seat, safety belt, and shoulder harness that are designed to provide the occupant protection provisions required in § 23.562 of this part. Other seat orientations must provide the same level of occupant protection as a forward-facing or aft-facing seat with a safety belt and shoulder harness, and provide the protection provisions of § 23.562 of this part.” The emergency landing ultimate inertia load factors have remained unchanged from Amendment 23-36 through Amendment 23-54, effective December 20, 2000. Amendment 23-54 is the latest amendment level to part 23. For inertia force requirements for occupant protection preceding CAR 3, refer to Table 1 in AC 21-34, which lists the requirements for the regulations dating from Bulletin 7-A to the original part 23. What are the methods of approval for retrofit shoulder harness installations? 1. Supplemental Type Certificate (STC). An STC is the most desirable and most rigorous approval. The STC offers the highest assurance that all of the airworthiness regulations have been met. The STC approvals are issued by the FAA Aircraft Certification Offices (ACO's). STC approvals are usually obtained ### Page 104 92 by a shoulder harness installation kit supplier for multiple airplane installations in an airplane model or model series. AC's 21-34 and 23-4 (References 1 and 2) provide guidance and acceptable means of compliance for shoulder harness and seat belt installations. AC 23-4 specifically addresses part 23 installations. These AC's are also applicable to installations in airplanes having a certification basis of predecessor regulations (for example, CAR 3). An applicant for an STC may use a salvaged airplane fuselage to substantiate the strength of the fuselage and the shoulder harness attachment fittings by structural tests, since the shoulder harness attachment structural test may damage an airworthy fuselage. It may be a problem that the available test airframe may be stronger than the lowest strength production airframe. This may be a problem in steel tube airframes. During many years of producing such airframes, various specification materials may have been used. For example, many CAR 3 (and predecessor regulations) airplanes were originally produced from 1025 steel tubing and later constructed from higher strength 4130 steel. In one case studied, two different specification 1025 steel tubes were used which may have an ultimate tensile strength (UTS) ranging from 55,000 to 79,000 pounds per square inch (psi). The UTS of 4130 steel is 90,000 to 95,000 psi. The test article should be representative of the lowest strength production airframe. This may be accomplished by a conformity inspection using the production drawings. The strength of materials of parts affected by the modification needs to be verified by the airframe manufacturer’s process and production records. The serial number of the test article needs to be verified. An alternative course of action would be to determine, by appropriate tests (for example, chemical analysis, hardness tests, strength tests), the strength of the parts of the test article affected by the modification. Follow with testing to a conservatively higher load that accounts for the difference in strengths of the test article and the lowest strength production article. Determination of the higher applied test load should take into account any uncertainty in the test(s) used to determine the strength of the material. Another alternate course of action may be to conduct the harness pull test on the available test airframe. The applicant may then substantiate the strength of other tubing specifications by a combination of test results and analysis. ### Page 105 93 AC 23-4 provides an acceptable means of compliance for static strength substantiation of attachment points for occupant restraint system installations. A test block is described to apply the 9.0-g forward inertia load. The safety belt installation alone is tested to 100 percent of the load. The shoulder and safety belt combined load is distributed 40 percent to the shoulder harness and 60 percent to the seat belt. In airplanes having side-by-side seats, the pull test may need to be applied simultaneously to the harness fittings for both seats. However, this depends on the type of harness and where the upper ends are anchored. Normally, this would not be necessary for a single diagonal belt shoulder harness attached to the outboard fuselage side or wing spar root end. In the case of a pull test for a retrofit shoulder harness installation in a tandem seated tubular steel fuselage, the forward inertia load was applied simultaneously for both harnesses. This was done for convenience in applying and reacting the loads. It was found, that due to the tube geometry, the load at the aft harness attachment caused a tension in the rear spar carrythrough tube. The front seat shoulder harness upper end was attached to the rear spar carrythrough tube. This enabled the front seat harness attachment to test to a higher load than if the pull test was done to each harness individually. In such a case, the test loads for each harness should be done individually. Part 21, § 21.50(b) requires the holder of an STC to furnish Instructions for Continued Airworthiness, prepared in accordance with § 23.1529. An STC can not be used to modify an aircraft without the permission of the STC holder. FAA Notice 8110.69 dated June 30, 1997, requires the STC holder to provide the customer (installer or airplane owner) with a signed permission statement that includes the following: - product (aircraft, engine, propeller, or appliance) to be altered, including serial number of the product; - the STC number; and - the person(s) who is being given consent to use the STC. The permission statement needs to be kept as part of the aircraft records. The requirement for this permission statement originated in the Federal Aviation Authorization Act of 1996 (Public Law 104-264). This provision was put into law to try to stop the pirating of STC's. ### Page 106 2. Field Approval. A shoulder harness installation in a small airplane may receive a Field Approval (FAA Form 337) granted by a Flight Standards Aviation Safety Inspector. Field Approvals are appropriate for alterations that involve little or no engineering. If the installation requires structural modifications, an Aircraft Certification Office will need to assist in the Field Approval process by approving the structural aspects of the installation. A Field Approval constitutes a change to type design and must meet the same regulatory requirements as an STC. AC 43.13-2A (Reference 3) contains methods, techniques, and practices acceptable to the Administrator for use in altering civil aircraft. Chapter 9 covers shoulder harness installations. Section 3 covers attachment methods. Shoulder harnesses installed under Field Approval must meet the same regulatory requirements as an STC; therefore, the applicant should demonstrate by test 9.0- g forward load capability. The test load should be 814 pounds for Normal Category or 910 pounds for Utility or Acrobatic Category, in accordance with AC 23-4. Reference 4, Chapter 1, Perform Field Approval of Major Repairs and Major Alterations, Section 1, paragraph 5D(2) states: “Acceptable data that may be used on an individual basis to obtain approval are: - AC's 43.13-1A and 43.13-2A, as amended* - Manufacturer’s technical information (for example, manuals, bulletins, kits, and so on) - FAA Field Approvals” * Note: Advisory Circular (AC) 43.13-1B, dated September 8, 1998 superseded AC 43.13-1A. When using a previous Field Approval as acceptable data, the pull test need not be done if it can be determined that a previous pull test applied 814 pounds for Normal Category or 910 pounds for Utility or Acrobatic Category. Field Approvals for shoulder harness installations should not be done by referencing a previous Field Approval and deleting the pull test, unless the attachment parts have a Parts Manufacturer Approval (PMA), or other FAA approval. If the attachment parts have no FAA approval, the strength is not known or assured, since they have not been manufactured to an FAA approved quality control system. Shoulder harness installations attaching to the center of an unsupported wing carrythrough tube, or other unsupported member, should not receive a Field ### Page 107 95 Approval without a design approval from an Aircraft Certification Office. Applying the test load in such cases may cause damage or permanent set to the affected structure. Existing FAA guidance, including AC 43.13-2A and AC 21-34, recommend against attachment to the center of unsupported members. Figure 9-16 in AC 43.13-2A shows typical shoulder harness attachments to tubular members. These are all at tube intersections and not at the center of unsupported tubes. Figure 9-12 shows a typical wing carrythrough member installation. This appears to be in the center of the carrythrough member that is a hat section as found in metal skinned airplanes. Part of the figure shows that the hat section is riveted to sheet metal skin (which would provide longitudinal support). Personnel performing the Field Approval must ensure that both the harness and belt are compatible and have a TSO approval. 3. Minor change. Part 21, § 21.93(a), Classification of changes in type design, states: “A minor change is one that has no appreciable effect on the weight, balance, structural strength, reliability, operational characteristics, or other characteristics affecting the airworthiness of the product.” Information provided to us by the Anchorage ACO indicates that some shoulder harness installations that provide known safety improvements have been approved as a minor change. In these situations, the FAA certificated mechanic who installs it makes an entry in the maintenance log of the airplane. One shoulder harness installation kit supplier uses this process (no FAA approvals) to install shoulder harnesses in PA-18 airplanes. The installation does not require modification of the airframe. The front seat harness attaches to the center of the rear wing spar carrythrough tube. However, it may not meet the 9.0-g forward inertia load required by CAR 3.386. The kit supplier stated that some airplane owners who had accidents reported that the harness installation had saved their lives. In general, shoulder harness installations should not use the center of an unsupported wing carrythrough tube or other unsupported member as an attachment point. This type of attachment may pose a risk to the structural integrity of the airplane. Although the attachment may be a clamp-on fitting that does not alter the existing airframe, the installation may result in a major change in the type design. This is because the shoulder harness attachment may introduce new loading conditions into the carrythrough tube. ### Page 108 96 It is acceptable for the carrythrough structure to be damaged in an emergency landing. However, it is unacceptable for the tube to fail in-flight. Carrythrough tubes, highly loaded in compression, may experience a beam-column buckling failure if the occupant applies a load to the shoulder harness attachment. In some cases, very small loads on the shoulder harness attachment may cause beamcolumn buckling failures. Some shoulder harnesses that have been installed by minor change do not have a TSO approval. TSO-C114, Torso Restraint Systems, was issued March 27, 1987. Torso restraint systems manufactured before that date did not have to meet the prescribed Society of Automotive Engineers standard, Aerospace Standard 8043, Aircraft Torso Restraint System, dated March 1986. AC 43.13- 2A and AC 21-34 provide guidance for acceptable harnesses. Acceptable harnesses for minor change installations include the following: - harnesses that meet TSO-C114 or Military Specification (MIL-SPEC) requirements, - harnesses that have been produced under a Parts Manufacturer Approval (PMA), or - other harnesses appropriate to the certification basis of the aircraft. We have studied the circumstances and legality of shoulder harness installations done by minor change. An airplane owner may wish to install shoulder harnesses, but an STC or prior Field Approval is not available for his airplane. In this case, it is not likely that an individual airplane owner would apply for an STC or a Field Approval since the owner would need to hire an engineering consultant to perform the structural test and any associated structural analysis. Also, there is a possibility that the airframe may be damaged during the pull test. In such installations, a pull test would not be done and there is no assurance that the installation will provide occupant protection to the ultimate inertia force requirements (particularly the 9.0-g forward force) of § 23.561 or CAR 3.386. Concerning the legality of shoulder harness installation by minor change, we conclude the following: Since CAR 3.386 and § 23.561(b)(1) before Amendment 23-36 (which became effective September 14, 1988) state that “proper use is made of belts or harnesses provided in the design,” the previously approved seat belt installation alone must meet the prescribed ultimate inertia forces. Civil Air Regulation (CAR) 3.652, Functional and installation requirements, states: “Each item of equipment which is essential to the safe operation of the ### Page 109 97 airplane shall be found by the Administrator to perform adequately the functions for which it is to be used, shall function properly when installed, and shall be adequately labeled as to its identification, function, operational limitations, or any combination of these, whichever is applicable.” Before Amendment 23-20 (which became effective September 1, 1977), § 23.1301 contained essentially the same requirement as CAR 3.652. Amendment 23-20 deleted the words “essential to safe operation” and made the provisions of § 23.1301 applicable to “each item of installed equipment.” We conclude from these rules that if a shoulder harness is not required equipment, it is not essential to the safe operation of the airplane; therefore, CAR 3.652 and § 23.1301, before Amendment 23-20, should not be used as a basis to prohibit shoulder harness installation by minor change. These rules should be applied to shoulder harness installations made by STC and Field Approval. The mechanic making such installations should consult AC 43.13-2A, Chapter 9, for information on restraint systems, effective restraint angles, attachment methods, and other details of installation. See AC 21-34, Shoulder Harness-Safety Installations, AC 21-25A, Approval of Modified Seats and Berths Initially Approved Under a Technical Standard Order, and AC 43.13- 2A, Acceptable Methods, Techniques, and Practices—Aircraft Alterations. Aft-facing and side-facing seats For aft-facing seats, seat obliqueness should be limited to 15° unless additional occupant protection for side-facing seats is installed. Part 23 permits side-facing seats, but it does not address the crashworthiness problems of these installations. We recommend that side-facing seat installations be discouraged. If such an installation is made, the following should be applied in addition to any applicable rules from the original certification basis: a. A sideward facing seat is defined as one in which the plane of symmetry of the occupant makes more than an 18° angle with the vertical plane containing the airplane centerline when viewed from above. b. Each occupant of a sideward facing seat should be protected from serious head injury when experiencing the inertia forces of § 23.561(b)(2) by either a safety belt and energy absorbing rest that will support the head and torso or by a safety belt and shoulder harness that will prevent the head from contacting any injurious object. There should be adequate padding on any restraining bulkhead. Riding up of diagonal ### Page 110 98 shoulder straps on the neck, which could cause neck injury, and location of attachments and rigidity of the seat support that could cause twisting and compression of the spine should be considered. For a multiple side-facing seat, a passenger seated immediately forward of another passenger cannot be considered an energy absorbing rest (human cushion). c. Sideward facing seat installations that do not comply with paragraph “b.” above should be placarded to prohibit occupancy during takeoff and landing. In any case, the side-facing seats still require one seat belt for each passenger to protect against inflight turbulence, and the berth should be considered an item of mass for emergency landing conditions of § 23.561. d. Special conditions will be required for sideward facing seats that are to be occupied during takeoff and landing when the certification basis requires compliance with § 23.562, Emergency landing dynamic conditions. Seat Removal for Parachute Operations This guidance is intended for engineering assistance in the field approval process, Type Certificate (TC) approval or Supplemental Type Certificate (STC) approval of sport parachute modifications to small airplanes. The Small Airplane Directorate has received National Transportation Safety Board (NTSB) recommendations regarding the safety record of small airplanes that have been modified for use in sport parachute operations. Most of these modifications have been approved by FAA Form 337 Field Approvals without engineering assistance. These modifications should either be approved by the TC process, STC process or by a field approval with engineering assistance. Aircraft Certification Offices (ACO's) should not provide engineering assistance for sport parachute modifications that change the limitations of an airplane or affect primary structure, required systems and equipment, handling, crashworthiness or performance. They should be completed using an STC or TC. The acceptability of sport parachute operations is given in operational regulations. 14 CFR part 91 allows the use of the cabin floor for sport parachute operations and requires the operations be conducted per 14 CFR part 105. Part 105 requires radio equipment for operations in controlled airspace, but there are no other equipment requirements other than the parachutes. Advisory Circular AC 105-2C, Sport Parachute Jumping, gives information on complying with the regulations and provides a list of airplanes that may be operated with one cabin door removed. The part 91 regulation means floor seating is acceptable only for sport parachute operations, but this does not mean that the crashworthiness of the airplane can be unacceptable or that an unsafe condition can be created by a modification. ### Page 111 99 The small airplane certification bases are mostly Civil Air Regulation (CAR) 3 or early part 23. The modifications include the following: seating capacity increases, seat and restraint removal and replacement belt installation, door removal for airplanes not listed in AC 105-2C, step and hand holds on airplane exteriors, gross weight increases and center of gravity (CG) changes. Seating capacity increases can be acceptable if all persons are protected in emergency landing loads per § 23.561 (3.386), all persons are secured against sliding backwards in climbs to maintain an acceptable CG position, and any changes in allowable CG range and increased gross weight are found acceptable. The applicant should also supply data to show that emergency exits are not crowded in a ground evacuation with the increased seating capacity per § 23.807 (3.387). This test/analysis must be done with all occupants except one representing a pilot in a parachute, and all parachutists initially restrained in place by the replacement belts/restraints. Removal of certified seats, seat belts and shoulder harnesses, if any, is allowable per part 91, § 91.207(a)(3)(ii). This does not remove the requirement to prevent an unsafe condition in an emergency landing or normal operations. The replacement belts/restraints should, therefore, meet the requirements for emergency landing upward, forward and sideward loads, per § 23.561 (3.386), and secure all persons against sliding backwards in climbs to maintain an acceptable CG position. Per § 23.785 (3.390), the attachment of the replacement belts/restraints to structure must have a factor of 1.33 times the emergency landing loads in § 23.561 (3.386). A simple belt that fastens across the lap of a person sitting on the cabin floor is not acceptable. This prevents movement only in the direction the person is facing. For instance, facing to the rear (as has been reported) prevents movement toward the tail in a climb, but it does not secure a person against uncontrolled movement toward the front in an emergency landing. The FAA and the United States Parachute Association (USPA) have tested a pair of adjustable belts attached to the parachute harness and found that they are more effective in providing restraint while the emergency-landing forces dissipate. We find it acceptable, therefore, for a parachutist to use adjustable belts instead of the seat belt. Acceptable restraint systems for floor seating of sport parachutists are given by the Civil Aeromedical Institute (CAMI) of the FAA in Report DOT/FAA/AM-98/11, Evaluation of Improved Restraint Systems for Sport Parachutists, dated March 1998. The dual-strap restraints (5D, 6D and 7D) provided the least flailing and bending of body segments and the least forward translation of the pelvis. ACO's should require one of these restraint systems for any TC, STC or engineering assistance to a field approval. Limitations should be imposed to call out the three operational procedures on Page 21 of the CAMI Report. The dual strap restraints have been tested with floor attachments. To date, there has not been any acceptable testing of a sidewall attachment. An applicant who wants to attach a restraint system to a sidewall will need to perform sled testing per an approved procedure to verify its suitability in the emergency landing loads. ### Page 112 100 An acceptable procedure for use of the dual-strap restraints (5D, 6D and 7D) is as follows: a. Sit on the aircraft floor between two seat rails facing aft. b. Attach two adjustable belts to your parachute harness per either the 5D, 6D or 7D configuration. You may loop the adjustable belt through your parachute harness webbing or you may attach a quick-release clip through the ring on your harness. c. Attach the other end of each harness to the seat rail in the aircraft floor. Make the attachment to the floor rail on the same side as the attachment to your harness. d. Adjust the belt as snugly as possible without causing discomfort. Note: If possible, place a lap belt attached to the floor rails around your upper legs to prevent flailing during the emergency landing. Door removal raises issues in Subpart B of 14 CFR part 23. Performance, stability, and control must be found acceptable with the door removed in flight. If the Airplane Flight Manual gives performance numbers such as takeoff and takeoff over a 50 foot obstacle, the applicant should verify that performance or issue a Flight Manual Supplement with the revised numbers. Of course, the applicant does not have to present any data or do any tests for airplanes that are listed in AC 105-2C. Steps and handholds to be installed on the exterior of a small airplane should be shown to meet the structural strength and fatigue requirements in part 23 (CAR 3). Carriage of parachutists outside the aircraft must be shown to not be hazardous in a power loss failure in takeoff and climb flight phases. Increases in gross weight and allowable CG range should be shown to comply with the requirements in part 23 (CAR 3) as would any other modification that affected these numbers. Aircraft modified for sport parachute operation must include a placard or flight manual limitation that prohibits flight with doors or seats removed except in sport parachute operation. Amendment 23-7 and Subsequent AC 23-28, when issued, will provide information and guidance applicable to the static strength substantiation of the attachment points for occupant restraint system installations, which have both a safety belt and shoulder harness. ### Page 113 101 Amendment 23-36 and Subsequent See guidance for § 23.562, Emergency landing dynamic conditions. Part 23 did not envision more than two seats on the flight deck, although the part does not prohibit such an installation. The airworthiness standards do not contain adequate standards for an “observer” seat (occupied by an FAA Flight Standards inspector on commuter flights). Therefore, we would expect to apply special conditions to such an installation that would address occupant restraint, emergency egress, and appropriate placarding to prohibit use by a passenger under any circumstances. The special conditions should establish a level of safety equivalent to that established in the certification basis of the airplane, not only for the observer seat occupant but also the crewmember seated in front of the occupant. The weight of a parachute is included in the 215-pound occupant weight if a parachute is required. The weight remains at 215 pounds when a parachute is not required. ### Page 114 23.787 Baggage and cargo compartments Original Issue and Subsequent Questions have been raised regarding § 23.787(c), which requires an ultimate inertia forward force of 4.5g for the protection of passengers from any cargo compartment. This regulation is related to Civil Air Regulations, which envisioned a crew compartment forward, a passenger compartment in the middle, followed by a bulkhead and a small cargo/baggage compartment aft. In this concept, the 4.5g was considered adequate based on NASA data that showed g forces become less as distance from the nose increases in a typical crash. Our review of all cargo configurations has led to the conclusion that under § 23.561(b) and (e), the restraining devices should meet the 9g requirements. The up and side load inertia forces are not considered to be applicable in this case where the crew would not be subject to injury from upward or sideward cargo movement. To modify a passenger plane to an all-cargo configuration, the following items should be considered: a. The cargo compartment should meet the requirements of § 23.787. Special attention should be given to cargo loading placards and the cargo restraint system. b. The cargo restraint system, including tie downs and the supporting structure to which they are attached, should be substantiated to the emergency landing ultimate inertia forces in § 23.561(b)(2). c. The floor loading should be re-substantiated to ensure the floor structure is not overloaded. d. Emergency egress from an emergency exit or the entrance door should be verified accessible for the crew. e. A supplement to the Airplane Flight Manual (AFM) weight and balance section that shows the various permissible cargo loading arrangements and cargo restraints should be furnished. Guidance on Carriage of Hazardous Cargo for Operators of part 23 Certificated Aircraft There are no 14 CFR part 23 airworthiness standards that directly address the design of an aircraft to allow for the carriage of hazardous cargo and passengers simultaneously. We cannot, therefore, provide guidance on the issuance of Amended Type Certificates (ATC’s) or Supplemental Type Certificates (STC’s) specific to this kind of operation. ### Page 115 103 However, for the carriage of cargo and baggage in general, 14 CFR part 23 does provide certification requirements for baggage and cargo compartments. Specifically, § 23.787 “Baggage and cargo compartments,” which focuses on the design requirements of the baggage or cargo compartment and its integration with the airframe. § 23.855 “Cargo and baggage compartment fire protection” focuses on the shielding, insulation, and flammability requirements of cargo or baggage containers. In addition, this regulation adds additional requirements and tests on cargo and baggage compartments for use in commuter category aircraft. Title 49, § 175.85, states that hazardous materials may be carried in a main deck cargo compartment of a passenger aircraft provided the compartment meets all certification requirements for a Class B aircraft cargo compartment as defined in 14 CFR part 25, § 25.857. The certification requirements for Class B cargo compartments exceed the requirements identified in §§ 23.787 and 23.855. Normally, a Class B cargo compartment is not a certification requirement for a part 23 certificated airplane. However, a cargo compartment may be certificated, using either an ATC or STC process possibly involving special conditions, if it meets the appropriate airworthiness design standards for a Class B aircraft cargo compartment. Any ATC/STC effort must be coordinated with the responsible Aircraft Certification Office (ACO). Once approved, a Class B cargo container installation in a part 23 certificated aircraft does not necessarily entitle the operator to carry hazardous cargo. The operator must also demonstrate compliance with the appropriate operational rules and requirements of 14 CFR part 135 and Title 49 governing the carriage of hazardous cargo. For additional information on part 135 requirements, we recommend the operator contact their local Flight Standards District Office (FSDO). Amendment 23-14 and Subsequent See AC 23-2, Flammability Tests. Amendment 23-36 and Subsequent The rigid moveable/removable cargo restraint bulkhead attached to seat rails and to points along the cabin sidewalls and roof is considered a structure per § 23.787(c). Prior to this amendment, the loads to design this structure were not defined, but the loads for a cargo restraint system and tie downs in a cargo compartment had to withstand an ultimate inertia force of 4.5g. Even though not defined, some certification programs applied a 4.5g ultimate load factor to design a rigid moveable/removable cargo restraint bulkhead in the ### Page 116 104 cabin. The rationale was to bring the sum of occupant protection to a 9g forward load. In this amendment, the ultimate forward load factor for any cargo restraint system and tie downs has been increased to 9g. In this case, the structure can be designed to no load since the occupant protection of 9g has been met by the cargo restraint system and tie downs. If this structure separates the occupant compartment from the cargo compartment, only § 23.787(c) applies. Section 23.787(b) is applicable if cargo is carried aft of the occupants in the same occupant compartment. The ultimate load factor in § 23.561 has been increased to 18g by Amendment 23-36. ### Page 117 23.791 Passenger information signs No policy available as of June 30, 2001. ### Page 118 23.803 Emergency evacuation Amendment 23-34 and Subsequent See AC 20-118A, Emergency Evacuation Demonstration, for information on how to conduct an emergency evacuation demonstration of a commuter category airplane. If there is a project for a litter installation for non-ambulatory passengers, then the airplane should be evaluated for compliance with the applicable egress requirements for those passengers who can exit the airplane under their own power per § 23.803. This evaluation can be a simple engineering judgment if it is clear the litter installation will not prevent the safe egress of all non-litter passengers within the allotted time. If there is doubt, a new demonstration should be run that evaluates the ability of non-litter passengers to exit the airplane with special attention to the litter installation and possible obstructions to safe exit. Amendment 23-46 and Subsequent This amendment adds a requirement for emergency lighting per § 23.812 to be the only lighting used in an emergency evacuation demonstration when certification of emergency exits is done per § 23.807(d)(4). AC 20-118A is still applicable with the exception of paragraph 5a(3)(vi). ### Page 119 23.805 Flight crew emergency exits Amendment 23-46 and Subsequent Section 23.805 requires two crew compartment exits or a singular hatch if the passenger exits are not convenient and readily accessible to the crew, while § 23.807(a)(3) requires a single crew compartment exit if there is a door that is likely to block the crew's access to the passenger exits. If a crew compartment exit is necessary under § 23.807(a)(3), the more stringent requirements of § 23.805 are also applicable. In other cases, the crew compartment exits may be required by § 23.805 independent of § 23.807(a)(3). Crew compartment exits that comply with the requirements of § 23.805 would also comply with § 23.807(a)(3) without further showing. Amendment 23-49 and Subsequent Crew compartment exits that comply with the requirements of § 23.805 would provide an equivalent level of safety to § 23.787(c). According to the preamble to Amendment 23- 49, which added § 23.787(c), “this requirement would provide increased assurance that flight crews of all cargo airplanes will have ready access to an emergency exit.” ### Page 120 23.807 Emergency exits Original Issue and Subsequent “Seating Capacity” as used in this regulation is defined as the number of occupants, both crew and passengers, for which the airplane is certificated. Consequently, removal of installed seats is not justification for removal of certificated emergency exits. The regulation requires a clear and unobstructed opening. An exemption per 14 CFR part 11 to § 23.807(b) is acceptable for a seatback that protrudes into the opening if it can be easily pushed forward to clear the exit without requiring an action to unlock/unlatch the seat. If a seatback clears the exit when upright but not when reclined, it is acceptable to placard the seat to be upright during takeoff and landing. Emergency exits should be located to allow escape without crowding in any probable crash attitude. The inverted position is considered probable for both tail wheel and tricycle gear airplanes. This applies to airplanes with doors, forward sliding canopies, rearward sliding canopies and jettisonable canopies. If escape in an inverted attitude is not obvious or is questionable, then compliance should be demonstrated. It is not acceptable for certification purposes, except for acrobatic airplanes (§ 23.807(b)(5)), to rely on an emergency procedure requiring canopy jettisoning before an accident occurs. Regarding the acrobatic category, if the canopy is not jettisonable, it should be shown that the canopy can be opened far enough in flight between VSO and VD to enable the occupants to safely exit the airplane. If jettisonable, it should be shown that the canopy trajectory will not cause injury to the occupants while separating from the airplane between VSO and VD. Also, if the canopy is jettisonable, it should be demonstrated that the airplane can be safely flown without the canopy, or that an inadvertent jettisoning is shown to be improbable. Regarding doors between the pilot’s compartment and the passenger compartment that are likely to block the pilot’s egress in a minor crash landing, there should be an exit in the pilot’s compartment. This does not apply to curtains suspended from a rod at the top and made of flexible material without slats on any side. Doors or folding doors with rigid-frangible materials may jam in a minor crash. Acceptance of frangible doors can be shown by the evacuation procedure in paragraph 23.807(a) below or by the conditions for acceptance of rigid doors in paragraph 23.807(b) below. Rigid doors are only acceptable by placarding the doors to be latched in the open position for takeoff and landing, providing the conditions in paragraph 23.807(b) below are in compliance. ### Page 121 a. The purpose of the test is to demonstrate that the door between the pilot’s compartment and the passenger compartment will not block the pilot’s escape in the event the door is jammed. Acceptable means of compliance is by demonstrating the door is frangible and the flight crew can egress the airplane without assistance within the 90-second time limit. (1) The test should be conducted in an airplane or a mockup that conforms to the production airplane interior configuration that contains a bulkhead and door to be tested. The door should be closed to simulate jamming. If fragments from the broken door could obstruct the escape route of passengers and an emergency evacuation demonstration is required by either airworthiness or operating rules, then consideration should be given to including passenger participants in the test. In this case, refer to § 23.803 for guidance. (2) Two participants representing a pilot and a copilot will be used in the test. They should be persons with no particular escape abilities. The approximate stature and weights for the participants should be a female 60 inches tall weighing 102 pounds and a male 74 inches tall weighing 210 pounds (5th to 95th percentile). The female participant will break the door and be the first person through the exit without assistance from the male participant. Instructions for enhancing the egress should be limited to those instructions that are provided in either the FAA approved Airplane Flight Manual (AFM) or on related placards, or both. (3) Determine that the lighting simulates night lighting with no moonlight or starlight. Lighting may be allowed at ground level to aid in leaving the area near the airplane providing the lighting is kept low and is shielded so it does not aid in evacuating the airplane. (4) Participating personnel should be informed of the purpose of the demonstration and of the safety precautions. Safety of participants is the responsibility of the applicant and safety procedures should protect the applicants without impacting the test results. Participants may wear protective gear such as helmets, but such gear, tools, or any other device should not be used to break through the door. (5) The time limit is 90 seconds whether or not passenger participants are used in this demonstration. (6) Information advising the flight crew that the door is frangible should be placarded on the door(s) and should be noted in the Limitations Section of the AFM. ### Page 122 b. Rigid doors (those with stiff members that may jam in a minor crash) may be approved providing they are placarded to be latched open during takeoff and landing and under the following conditions: (1) The opening and latching should be included in the Normal Procedures Section under the Before Takeoff and Before Landing Checklists of the AFM. (2) With the door latch in the fully open position, the latch should be able to withstand the loads from the ultimate forces relative to the surrounding structure, per § 23.561. (3) Flight crew members should be able to open and latch the door with their safety belts/shoulder harnesses fastened, if required by either airworthiness or operating rules. (4) If certification for night operation is requested, the pilot’s compartment—with the doors open—should be free from glare and reflections that could interfere with a pilot’s vision, per § 23.733. (5) The doors should be placarded in accordance with § 23.1557. The placards should state that the doors are to be latched in an open position before takeoff and landing. Placard information should be in the Limitations Section of the AFM. The following should be considered when approving supplemental type certificates for cargo conversions (no passenger seating) on airplanes with a certification basis of 14 CFR part 23, including commuter category airplanes: • Unless otherwise specified, the airplane “seating capacity” is defined as the number of occupants, both crew and passenger, for which the airplane is certificated. Simply removing seats from the seat rails does not change the seating capacity. • The § 23.807(a)(3) conditional requirement for one crew compartment exit is applicable to all airplanes with a part 23 certification basis, including commuter category. • Airplanes with a crew compartment door that may block the pilot’s exit must have an exit in the crew compartment. In this circumstance, a normal, utility, or aerobatic category airplane with a certificated passenger seating capacity of zero (cargo only), a minimum of a single accessible exit in the cargo compartment in addition to the crew compartment exit is necessary (reference § 23.807(a)(3)). ### Page 123 • It is not acceptable to eliminate emergency exits by modifying an existing exit so it no longer complies with 14 CFR part 23 airworthiness requirements. However, under certain circumstances, such as in an airplane with more than two exits, it may be permissible to block exits with cargo provided one or more exits remain accessible on each side of the airplane. Implicit within the definition of “readily accessible” is a requirement that the unblocked emergency exits have no obstacles such as cargo nets or cargo. Obstacles pose an impediment to the emergency evacuation process that cannot be quantified, particularly when consideration is given to the effect of a post-crash environment. An emergency evacuation demonstration, therefore, is not acceptable as an equivalent level of safety since it is not meant to allow alleviation or deviation of specific requirements. If the airplane is configured for both passenger and cargo (Combi), the requirements of this section should be met for the passenger compartment. In addition, cargo should be located so that it does not obstruct either access to or use of any required emergency or regular exit; so that it does not obstruct the use of the aisle between the crew and passenger compartment; and so that it meets the additional requirements of § 135.87. Amendment 23-10 and Subsequent Multiseat airplanes must have a second exit on the opposite side from the main door, per § 23.807(a)(1). Both exits must be accessible and must not be blocked with cargo. Amendment 23-34 and Subsequent This amendment added emergency exit requirements for commuter airplanes. Included are emergency exit marking requirements as well as those in §§ 23.783 and 23.1557. The additional emergency exit marking requirements in § 23.1557(d) regarding the red operating handle and placard that provides door opening instructions are not mandatory for the passenger entrance door. However, § 23.807(b)(3) requires markings for easy location and operation of the exit even in darkness, and § 23.811(b) requires the illumination of the exit sign. As an added safety feature, it is recommended that the operating handle be self-illuminated and marked with a red arrow and the word “OPEN” in red letters placed near the head of the arrow. If necessary, other pertinent instructions for opening the door should also be in red. This amendment requires three emergency exits as well as the entrance door for commuter airplanes with passenger seating from 16 to 19. Part 25, § 25.807, requires two Type III emergency exits on opposite sides of the cabin. It is possible for an applicant to use part 25 for emergency exits. To do so requires a petition for exemption per part 11, and compliance to part 25, §§ 25.807, 25.561(b)(3)(iv), 25.783, 25.809, 25.811, 25.812, 25.813, 25.815, and 25.817. ### Page 124 112 An integral stair, if installed at an entrance door, should be designed so it does not reduce the effectiveness of the door as an emergency exit under the inertia forces of § 23.561 and following the collapse of one or more legs of the landing gear. An actual demonstration of this failure mode is beyond the intent of this rule. It should be shown by orthographic drafting techniques or test (i.e., ground plane under an airplane to simulate various attitudes) that with the various combinations of collapsed landing gears and resulting airplane attitudes, the exit effectiveness is not reduced. This is done with no fuselage deformation. There are no standards for ejection seats in part 23. If an applicant needs an ejection seat to meet the emergency exit requirement in § 23.807, an equivalent level of safety will have to be justified. Emergency Exit Size and Shape Background For a commuter category airplane with a certificated passenger seating capacity of zero (cargo only), two emergency exits (one on each side) are required in addition to the main door (reference § 23.807(d)(1)(i)) regardless of the crew compartment door configuration. CAR 3.387 and 14 CFR part 23, § 23.807, have required that all emergency exits have sufficient size and shape to admit a 19 x 26 inch ellipse. Time to egress through an exit is related to the total open area and the most critical dimension of the exit. The area of a 19 x 26 inch ellipse is 388 square inches. Studies for emergency evacuation demonstrations with the standard ellipse have shown that the duration to egress was equal or less with other exits having a total open area equal to or greater than 388 square inches and the most critical dimension, width or height, greater than 19 inches, but lacking the shape to admit a 19 x 26 inch ellipse. Acceptable Means of Compliance Alternatives for compliance to the airworthiness standards are permitted by an equivalent level of safety. One method for determining compliance by an equivalent level of safety is by the test procedure below. Demonstrations have shown that the emergency exit size and shape greatly affect the time and ease of an emergency evacuation. An equivalent level of safety should only be considered if the exit meets the logical limits which correspond to the standard exit; that is, the total open area is equal or greater than 388 square inches and the most critical dimension, width or height, is not less than 19 inches. These limits for area, width, and height were established after considering human factors, evacuation demonstrations, and existing airworthiness standards. ### Page 125 113 Test Procedure Area of opening. The following factors should be considered when measuring or computing the area of opening: a. Firm protrusions that would hamper egress should be eliminated from the minimum required exit opening. Examples are seals or escape latches that will not easily compress, move, or fold out of the opening with the motion of a person moving through the opening. b. When a compressible seal protrudes into an opening, the seal may be in the compressed condition when measuring or computing the opening area. c. During the comparison test, the emergency exit opening used as a standard is an opening that will allow passage of a 19 x 26 inch ellipse with a major axis being in a vertical position, a horizontal position, or any other position. d. The area leading to the opening should be clear and unobstructed. Minor obstructions could be acceptable if there are compensating factors to maintain the effectiveness of the exit; that is, a total effective open area of 388 square inches and the most critical dimension, width or height, not less than 19 inches. For example, soft seatback cushions may constitute minor obstructions if the cushion can be readily moved away from the exit and the exit can be easily opened, and if the cushion in its normal position does not prevent identification of the exit or obscure the exit marking. Comparison Test Conditions. The comparison test will determine the difference in mean escape time between the proposed and standard exit or exits. a. A mockup of a section of the fuselage may be used. The arrangement of exits, passenger seats, and the step-up and step-down distances from the sill to the wing or step may be simulated. Ramps or stands are permitted to assist participants in descending from a wing when over-wing exits are used if the acceptance rate of the ramp or stand is no greater than that of the assist means of the airplane in an actual crash landing situation. Mats may be used on the floor or ground to protect participants. No other equipment that is not part of the airplane’s emergency evacuation equipment may be used to aid the participants in reaching the ground. b. At the start of each trial, participants should be seated as called out in AC 20-118A, Emergency Evacuation Demonstration. c. Participants should not be permitted any “practice” runs, but they may be briefed on the purpose of the test to demonstrate a rapid emergency evacuation of the airplane. They should not be briefed that the test is to compare exits. An example of an acceptable ### Page 126 114 instruction would be to pass through one foot first, followed by the head and the other foot. The briefing should be the same for each trial. d. The test should be conducted under dark or simulated dark conditions for both standard and proposed exit configurations per the Compliance Inspection Requirements of AC 20-118A. e. The Participant Composition should be as specified in AC 20-118A. Statistical Design. An acceptable statistical design is as follows: a. There should be 15 or more participants for each exit configuration to be tested, including the standard configuration. b. The participants should be assigned to the number of subgroups corresponding to the number of exit configurations to be tested. As noted in paragraph a., each subgroup should have at least 15 people unless the seating configuration is less than 15. In this case the following procedures should be used: (1) The subgroups should be divided into sub-subgroups of approximately equivalent size where the sub-subgroup size is equal to or less than the seating capacity of the airplane. The egress time of the sub-subgroups is totaled to constitute the subgroup time. (2) When a mockup for an airplane is used, even if the number of passenger seats is less than 15, the total subgroup of 15 participants may participate at the same time providing the increase of space from the standard mockup for the additional subjects does not degrade the comparison tests. Under these conditions, the participants with the least physical agility should be in the most critical positions. c. The subgroups should be as neatly alike as possible with respect to physical agility, age, sex, and weight. This can be achieved by first dividing the group by age and sex then subdividing each age/sex group at random into the required number of subgroups. d. Each subgroup should test each configuration, but the order of trials should be different for each subgroup as well as chosen in accordance with the Latin Square Principle. This principle is that each configuration be tried once by each subgroup and appear once in each possible order. Thus if there are two configurations to be tested and, therefore, two subgroups A and B, then Subgroup A should first try the standard configuration followed by the proposed configuration; Subgroup B should perform the trials in the reverse order. This arrangement eliminates the effects of an individual’s learning, fatigue, and agility. ### Page 127 115 Recording of Trials. Recording should be done as follows: a. Motion pictures or video recordings, sound or silent, should be made to analyze the trials for difficulties with an exit, individual escape times, and other performance factors. b. A large clock with a second hand should be placed in the camera field so that time can be recorded or synchronized electric cameras may be used with the time superimposed in the film processing. A signal light to indicate the beginning and end of each trial should also be arranged in the field of view of the camera. c. Evacuation time should be rounded to the nearest second. The timed demonstration is performed per the Evacuation Section of AC 20-118A. Evaluation of Results. The evaluation should be performed as follows: a. The effectiveness of the proposed exit or exits compared with the standard exit or exits is determined by comparison of the average time of the subgroups to pass through each exit tested. The effect of subgroup learning is canceled by the Latin Square Principle. b. It is possible that one group may contain one or two persons who find it difficult to go through the exits. The Latin Square Principle will cancel such unbalance between subgroups. c. It may happen that an individual may, through chance, have considerable difficulty with an exit, but their performance may compare with average performance of other individuals. A study of the individual escape times will enable such occurrences to be evaluated and will assist in the final determination of the acceptability of the proposed exit or exits. d. A proposed exit configuration is acceptable when its egress time is equal to or less than the time required to pass through the standard exit. Type Certificate Data Sheet. An equivalent level of safety should be part of the type certification basis and noted on the type certificate data sheet. Suggested wording is, “Equivalent Safety Findings: Section 3.387 of the CAR and Section 23.807 of 14 CFR part 23, emergency (particular) exit in accordance with AC 23-17, Systems and Equipment Guide for Certification of Part 23 Airplanes.” ### Page 128 23.811 Emergency exit marking Amendment 23-36 and Subsequent For small airplanes with emergency exits openable from the outside, the FAA recommends that markings be added to the outside of all exits as follows: a. Outline the exit with a band of a contrasting color from the surrounding fuselage surface. b. Mark the corners of the exit in a conspicuous manner. c. Outline the exit handle with a band of a contrasting color. d. Mark the exit with any other conspicuous visual identification scheme. e. Install a decal on the outside surface of the exit or the surrounding surface adjacent to the exit that shows the means of opening the exit, including any special instructions if applicable. Section 23.811(a) is specifically intended to be applicable to exterior emergency exit markings and placards. It does not stipulate the color of these markings and placards, with the exception that they must be conspicuously identifiable from outside the airplane. Section 23.1557(d) specifically requires that each placard and operating control for each exit must be red. Furthermore, a placard must be near each emergency exit control and must clearly indicate the location of that exit and its method of operation. Since the emergency exit exterior markings are conspicuously identifiable from the outside of the airplane, to further require red placards that indicate the location of these exits, does not make sense. There might be instances where the red placard may not be conspicuously identifiable due to the exterior color of the fuselage. Section 23.807 gives further guidance regarding adding external markings to emergency exits. Not one of the recommendations specifies the use of red placards. Other subparagraphs of § 23.811 require the use of red markings when the subject involves the interior of the airplane. Being consistent with these facts, therefore, we interpret the intent of § 23.1557(d) to be applicable to the interior of the airplane. Passenger exit signs should have an initial luminescence of at least 160 microlamberts, and should be replaced when its luminescence decreases below 100 microlamberts. ### Page 129 23.812 Emergency lighting No policy available as of June 30, 2001. ### Page 130 23.813 Emergency exit access See § 23.807 Means of Compliance for policy on acceptance of seatbacks obstructing emergency exits that can be easily moved. ### Page 131 23.815 Width of aisle Amendment 23-34 and Subsequent This rule applies to commuter category airplanes. The main passenger aisle width is the minimum distance between seats measured without occupants. This distance is measured without compressing the seat fabric or cushions, and with the seats and other aisle constraints in their most adverse position. If the seats can swivel and the distances can be less than the rule requires, the most adverse position may be defined with placards or flight manual limitations to require locking them in either a forward or aft facing position for taxi, takeoff, and landing. This is acceptable only when those positions give an aisle distance that meets or exceeds the requirement. ### Page 132 23.831 Ventilation Original Issue and Subsequent The use of an alternate air supply, either automatic or manual, that picks up air from within the engine compartment is unacceptable for cabin ventilation because of possible contamination from fuel, oil, or exhaust leaks. Halon 1301 may be safely used in concentrations up to 10 percent in airplane cabins. Ventilation in airplane cabins is sufficient for the agent to disburse in less than 5 minutes, so the time limit need not be considered if the concentration is held below the 10 percent limit. Halon 1211, however, should not be used in airplane cabins. The regulation regarding allowable CO concentration is given in part 23, § 23.831(a): carbon monoxide concentration may not exceed one part in 20,000 parts of air. Tests to verify compliance with part 23, § 23.831(a) are done as part of the original type certification of an aircraft when a modification that has a possible effect on the compliance to part 23, § 23.831(a) is made and when CO is verified or suspected in operation. Two acceptable methods of detecting carbon monoxide are given in Advisory Circular (AC) No. 20-32B. These are acceptable for a Type Certificate applicant, a Supplemental Type Certificate applicant, a Parts Manufacturer Approval applicant, an FAA Field Approval applicant, or an owner/operator. Other methods of verifying compliance to part 23, § 23.831(a) can be used if they are shown to provide acceptable results. Amendment 23-34 and Subsequent For pressurized airplanes, if hazardous accumulations of smoke are found to be reasonably probable in the cockpit area, smoke evacuation to a non-hazardous level should be readily accomplished from full pressurization to minimum safe levels (per 14 CFR part 91, § 91.211). Smoke evacuation procedures should be included in the Airplane Flight Manual, Emergency or Non-Normal (Abnormal) Procedures Section, or on approved placards. ### Page 133 **PRESSURIZATION** 23.841 Pressurized cabins Original Issue and Subsequent Paragraph (c) in § 23.841 requires there be a means to rapidly equalize the pressure differential. Assuming isothermal conditions, the time for the pressures to equalize depends on the cabin volume, the effective area of the safety-dump valves, the cabin inflow, and the pressures inside and outside the cabin. If the size of the effective area of the valve is small in comparison to the cabin volume, the rate of pressure change may be too slow to equalize the pressures before an adverse event could occur. The time period to rapidly equalize the pressures should consider maximum certificated cabin pressure differential, operation of the pressurization system, and either operation of the emergency exits or the cabin entrance doors, or both. When landing the airplane under emergency conditions, the safety-dump valve should have sufficient flow capacity to rapidly equalize the cabin pressure within a time period so that the cabin doors and emergency exits can be opened and evacuation is not impaired. Time to equalize the ambient and cabin pressures should be demonstrated. Paragraph (f) of § 23.841 requires a warning device for safe or preset pressure differential and absolute cabin pressure. A warning is interpreted to convey the need for an immediate corrective action, so it may not operate unless there is a failure, and the visual indication should be red per § 23.1322. Red lines on altimeters or pressure indicators are used to indicate operating limits, but they are not acceptable warning means. Inflatable door seals, if installed, are subject to the requirements of this rule. Amendment 23-14 and Subsequent This amendment requires that cabin pressure altitude not exceed 15,000 feet in any probable failure for airplanes certificated to operate over 31,000 feet. It is not appropriate to use an emergency descent procedure to demonstrate compliance to this rule when compliance can be achieved through design. The Airplane Flight Manual, Emergency Operations Section, should include an emergency descent procedure for use in a rapid decompression from any failure not withstanding the probability of its occurrence. Amendment 23-17 and Subsequent This amendment established 10,000 feet as the maximum absolute cabin pressure for operation of the pressure altitude warning. Therefore, the pressure sensors used in the warning system cannot have an operating set point and tolerance that would prevent the warning from being given at or before 10,000 feet. A feature that automatically changes the ### Page 134 122 warning altitude to 15,000 feet for operations at field elevations above 10,000 feet is acceptable to prevent nuisance warnings. The following material is a means of compliance to § 23.841(b)(3) that requires a means by which the pressure differential can be rapidly equalized. Section 23.841(b)(6) offers a provision for a warning indication at the pilot station to indicate when a cabin pressure altitude of 10,000 feet is exceeded. 1. RELATED 14 CFR PART 23 SECTIONS. These acceptable means of compliance refer to certain provisions of part 23 and the corresponding provisions of part 3 of Civil Air Regulations (CAR) in the case of airplanes for which those regulations are applicable. Listed below are the applicable and the related part 23 sections with the corresponding CAR sections shown in parenthesis: a. § 23.365 (3.197) b. § 23.775(c) (3.383) c. § 23.841 (3.395) d. § 23.843 (3.396) 2. DISCUSSION OF REQUIREMENTS. In discussing these requirements, a brief history on the development of the applicable airworthiness regulations is first presented. The purpose of the airworthiness requirements for small airplanes is then explained. a. Rapidly Equalizing the Pressure (1) History. The requirement for a means by which the pressure differential can be rapidly equalized was introduced in the airworthiness regulations for pressurized cabins for transport category airplanes when part 04 of the CAR became effective on November 9, 1945. Due to the trend to develop pressurized cabins for small airplanes, the 1956 Annual Airworthiness Review established similar requirements for pressurized cabins for small airplanes. The criteria were developed by using the principles that were applicable to pressurized cabins on transport category airplanes since most of the cabin pressure control system design for small airplanes drew heavily upon the equipment designed and developed for transport category airplanes. As a result, many of the provisions added to part 3 of the CAR by Amendment 3-2, effective August 12, 1957, were substantially the same as those which applied to transport category airplanes. Under the recodification program in 1965, part 23 replaced part 3 of the CAR and these requirements are now in § 23.841(b)(3). ### Page 135 (2) The purpose of this requirement is to provide the crew with a means to rapidly equalize the differential pressure to permit quick opening of the emergency exits and entry door(s) in the event of a gear up landing under emergency conditions. This means may be used for other events such as over pressurization and reducing cabin contamination. These functions are described in further detail as follows: (i) Due to a malfunction in the pressurization system or abnormal operational conditions, the cabin pressure is above normal conditions during the airplane landing phase. In this case, the cabin pressure may be vented by the safetydump valve operated through a manual controller or triggered by the landing gear safety switch so the emergency exits and the cabin entrance doors could be opened. (ii) If a failure such as a cracked window or windshield occurs, the cabin pressure should be capable of being rapidly reduced so the loads due to cabin pressure differential can be reduced accordingly. (iii) When a threatening cabin overpressure condition exists due to cabin pressurization system malfunction, the cabin pressure can be reduced by the safety-dump valve to prevent a structural failure of the pressure vessel. (iv) When the cabin air becomes contaminated by smoke, fumes, etc., the cabin safety-dump valve may be used, depending on the conditions, to assist the pressurization or ventilation system, or both, in evacuation of the cabin air to reduce the contaminants. b. Cabin Pressure Altitude Warning (1) History (i) The cabin altitude warning and many of the provisions for pressurized cabins for small airplanes were added to part 3 of the CAR by Amendment 3-2, effective August 12, 1957. Section 3.395(f) of part 3 of the CAR required, in pertinent part, that the pilot be provided a warning when safe or preset limits on pressure differential and on absolute cabin pressure were exceeded. (ii) In May 1958, a quantitative requirement was introduced in the airworthiness regulations when FAA established policy for altitude warning on the sport category airplanes. This policy, which was set forth in § 4b.375-l of CAR part 3, required that the warning for cabin pressure would meet the applicable requirements if it occurred when cabin absolute ### Page 136 124 pressure was reduced below that equivalent to 10,000 feet. Under the recodification program in 1965, part 25 replaced part 4b of the CAR and the 10,000 feet warning policy was carried over as an appropriate means of meeting the warning requirements in § 25.841. (iii) As part of the First Biennial Airworthiness Review Program in 1975, Amendments 23-17 and 25-28, which changed parts 23 and 25 respectively, were issued and became effective February 1, 1977. Amendment 25-28 transmitted a minor change to § 25.841 as follows: It changed "cabin absolute pressure is below that equivalent to 10,000 feet" to "cabin pressure altitude exceeds 10,000 feet." Amendment 23-17 brought into § 23.841 of part 23 a warning indication when the cabin pressure altitude of 10,000 feet mean sea level (MSL) is exceeded. The preamble for this change indicated this proposal was adopted because a large number of small airplanes had such a warning and many pilots had begun to rely on this warning. (2) The purpose of the cabin pressure altitude warning requirement is to indicate a warning at the pilot station when the cabin pressure altitude is greater than 10,000 feet MSL. A possible hazardous condition could be when the airplane reaches the operating altitude, which is greater than 10,000 feet MSL, and a malfunction in the cabin pressurization system occurs. If there was no warning for cabin pressure altitude, the cabin pressure altitude could slowly increase undetected to the airplane altitude, and the crew and passengers could become unconscious due to hypoxia. The effects of hypoxia are usually encountered when the flight crew is exposed to altitudes above 10,000 feet during extended flights. 3. ACCEPTABLE MEANS OF COMPLIANCE Warnings and Cautions. Section 23.1322 provides specific requirements for the assignment of red and amber for visual indications. Specifically, for abnormal operational or airplane systems conditions, a "caution" should be generated for crew awareness and subsequent crew action may be required; the associated color is amber. Under emergency operational or airplane systems conditions, a “warning” should be generated for immediate crew recognition and when corrective or compensatory action may be required; the associated color is red. If the cabin pressure altitude warning is a visual indicator, it should be red to indicate a hazard. ### Page 137 23.843 Pressurization tests Original Issue and Subsequent This rule applies to all doors. This includes doors that open outward, doors that open inward, and emergency exits. The 1.5 safety factor in § 23.303 does not apply when executing the pressurization tests in this section. Paragraph (a) of the rule specifies the pressure differential of § 23.365(d), which is 1.33 times the maximum relief valve setting. Paragraph (b) requires functional testing to verify operation so there cannot be any gross plastic deformation from the 1.33 factor. ### Page 138 **FIRE PROTECTION** 23.851 Fire extinguishers Amendment 23-34 and Subsequent See AC 20-42C, Hand Fire Extinguishers for Use in Aircraft, for guidance. ### Page 139 23.853 Passenger and crew compartment interiors Original Issue and Subsequent See AC 23-2, Flammability Tests. The purpose of this policy is for standardization in the approval of flammability testing of materials used in small airplanes per 14 CFR part 23, §§ 23.853, 23.855 and 23.1359. The Small Airplane Directorate policy for all certification projects, including engineering assistance for major alterations on FAA Form 337 Field Approvals, is that: a. The only materials for which flammability testing is not required are those that would not contribute significantly to the propagation of a fire per 14 CFR part 23, Section 23.853. These materials/parts are typically small parts. Material certification by an applicant or their supplier can be used in the determination that the material will not significantly contribute to the propagation of the fire. Company/Supplier material certifications cannot be used in lieu of official FAA flammability testing. b. All other proposed materials must be tested to the flammability level required by the certification basis and category of airplane. Advisory Circular (AC) 23-2, Flammability Tests, gives guidance on conducting flash-resistant, flame-resistant, fire-resistant, fireproof and self-extinguishing tests. The AC specifies that Appendix F of part 23 is to be used for proposed self-extinguishing materials. These are official FAA certification tests, which require witnessing by either an FAA certification engineer or an FAA Flammability Designated Engineering Representative (DER) who has authorization to witness a test. We will also accept from a DER submittal of an 8110-3 that attests to the validity of the data being approved. This can occur when the people running the test are well known to the DER and judged by the DER to be technically competent and reliable. Test data may only be approved by the FAA or by a DER with that approval authority. Note: DOT/FAA/CT-89/15” Aircraft Material Fire Test Handbook” is good reference material. c. A DER should not use FAA Form 8110-3 for flammability test results for a material when the testing is for quality assurance purposes for either a manufacturer or a repair station. Testing done for these purposes should be documented in a quality assurance report. When the material supports a certification project or an alteration or repair, FAA form 8110-3 is the DER’s only means of approving the technical data. The DER should determine if the testing documented in the test report adequately addresses the applicable airworthiness standards for the intended use of the material. If found ### Page 140 128 acceptable, the DER may generate an 8110-3 that references the test report. For the purposes of flammability testing, this technical data includes records of preconditioning of the test specimen. The flammability testing required by the certification basis is as follows: Passenger and Crew Compartment Interiors, § 23.853 Certification Basis Category Material Flammability Testing CAR 3 NormalFlame resistant if smoking allowed (1945-1946) Restricted Effective Experimental November 13, 1945 CAR 3 Normal, Flash resistant (1946-1949) Utility & Effective Acrobatic December 15, 1946 CAR 3 Normal, Flash resistant or flame resistant if (1949-1965) Utility & smoking is allowed in a specific Effective Acrobatic compartment November 1, 1949 Part 23 Normal, Flame resistant Through Amdt. 23-22Utility & (1965-1978) Acrobatic Effective December 5, 1978 Part 23 (78-present) Normal, Flame resistant or self-extinguishing Amdt. 23-23 & Subs Utility & per Appendix F for materials located & Subsequent Acrobatic on the cabin side of the firewall (1978-present) Effective December 1, 1978 Part 23 Commuter Self-extinguishing per Appendix F Amdt. 23-34 except for small parts in & Subsequent § 23.853(d)(3)(v) (1987-present) Effective February 17, 1987 ### Page 141 129 Cargo and Baggage Compartment Fire Protection, § 23.855 Part 23 Normal, Meet provisions of § 23.853(d)(3) Amdt. 23-49 Utility & (1996-present) Acrobatic Effective March 11, 1996 Part 23 Commuter Meet provisions of § 23.853(d)(3) Amdt. 23-49 (1996-present) Effective March 11, 1996 Electrical System Fire Protection, § 23.1359 Part 23 (96-present) Normal, Meet provisions of § 23.863 and Amdt. 23-49 Utility, § 23.1182 and wire insulation self Effective Acrobatic & extinguishing per Appendix F, March 11, 1996 Commuter 60 degree angle test 1. Per AC 23-2, cloth, wire and sheet specimens may be taken from a sample segment (batch/roll/sheet). In this case, FAA conformed test specimens/parts per 14 CFR part 21, § 21.33, are not required. Instead, conformity can be established for these types of materials on the basis of bill of materials, roll identification, etc. The 8110-3 should state it applies to the specific batch/roll/sheet for which the test was conducted. Note: Wire specified in AC 43-13-1B, Section 7, has been determined to be acceptable for use in certified airplanes and may be used without flammability testing. 2. In other cases, test specimens must be fabricated to accurately represent the production assembly or must be cut from actual parts. These parts should be conformed per 14 CFR part 21, § 21.33, prior to testing. 3. In both 1 and 2 above, a DER must comply with Order 8110.37C, Designated Engineering Representative Guidance Handbook. While we understand that in the past some DER’s have not always submitted 8110-3 forms with Make and Model information, 8110-3 forms must always be complete, including the Make and Model information. d. Flammability requirements have not been applied to conventional aircraft structure. However, the use of composite structure can result in a need to test a representative ### Page 142 130 build-up panel with an interior material, adhesive and composite structure, unless it is demonstrated the interior material does not permit an ignition source to penetrate it. e. Interior flammability tests may be required with build-up samples. Experience has shown that the thin exposed layer can burn away and expose the adhesive layer, which in many cases is extremely flammable and would contribute significantly to the propagation of a fire. Testing only the exposed layer without the adhesive backing would not be representative. Adhesives with a flame-retardant additive should be encouraged and listed; known flammable adhesives should not be used. f. We have reviewed the test criteria of part 25, Appendix F, part I and have determined that parts/materials tested to the part 25 test criteria are acceptable data to show compliance with the flame-resistant material requirement of § 23.853(a). ### Page 143 23.855 Cargo and baggage compartment fire protection See guidance for § 23.853. All-Cargo or Combination Passenger/Cargo Operations Special conditions will be required for airplanes approved for all-cargo or combination passenger/cargo operations because 14 CFR part 23 does not have applicable airworthiness standards for these operations. ### Page 144 23.859 Combustion heater fire protection No policy available as of June 30, 2001. ### Page 145 23.863 Flammable fluid fire protection Amendment 23-23 and Subsequent Those areas where ignition sources and flammable fluids are present must be designated as flammable fluid zones, and measures must be taken to minimize the probability of ignition sources and leakage being present. This is followed by a means to minimize the effects once a fire has occurred. This could be accomplished by extinguishing, ventilation, isolation, drainage, etc. The rule does not go so far as to make the entire airplane a “designated fire zone.” Fire detection and extinguishing is believed to be impractical for many part 23 airplanes. Where fire detection and extinguishing means might be impractical, the back up provisions could, for example, consist of a means to limit fluid leakage and fireproofing or isolation of critical parts. Therefore, compliance with § 23.863 could be accomplished with a means to limit fluid leakage, minimizing the probability of ignition, fireproofing or isolating critical parts. Minimization of ignition requires that equipment where a single failure can cause flammable fluid leakage, be tested to the explosion proof standards in RTCA/DO-160. If a finding is made that flammable fluids or vapors cannot escape into an area containing a potential ignition source or if the fluids are nonflammable, this rule would not apply to that area. Design measures could support the finding such as (1) shrouding (sealing off) of all potential ignition sources; or (2) shrouding or sealing off of all flammable fluid/vapor sources. In either case, it should be ascertained that the means will continue to serve its function following any single failure of the system or component it is isolating from the area. ### Page 146 23.865 Fire protection of flight controls, engine mounts, and other flight structure Amendment 23-14 and Subsequent The engine mounts refer to the aircraft structure for mounting the engine and not the mount pads or attachment points, which are integral parts of the engine. The intent of the regulations regarding engine mounts is that the engine remains in place with a fire heating an engine mount. We do not intend to cover the case of a general conflagration where the entire engine compartment is burning. Therefore, an applicant should design sufficient load paths for the engine to remain in place with a localized fire. For purposes of this rule, landing gears are not considered to be flight structures, so fireproofing or shielding landing gears is at the option of the manufacturer. Shielding made from fireproof materials in part 23, § 23.1191(h), may be used without flame testing. While the shielding may be made of fireproof materials that don’t require testing, means of installing the shielding such as sealers, adhesives, etc. should be shown to not reduce the efficacy of the shielding. Shielding materials subject to corrosion should be appropriately protected. Shielding need not be fireproof if it protects the enclosed structure to an extent equivalent to the enclosed structure being fireproof by itself. The effectiveness of such shielding or fireproof materials should be determined by subjecting the shielded or fireproof structure, or control, to flammability testing as defined in AC 23-2, Flammability Tests. Before removal of the flame at the end of the test, loads should be applied to the shielded structure or control to demonstrate that it can withstand the loads expected to occur during completion of the flight. These loads can be treated as ultimate loads. Amendment 23-48 and Subsequent Policy on Compliance of 14 CFR part 23, § 23.865 at Amendment 23-48 for Structures in Adjacent Areas Subjected to Effects of Fire in Designated Fire Zones The revision of 14 CFR part 23, § 23.865 at Amendment 23-48 includes: - changing the words “engine compartment” to “designated fire zones” for consistency with §§ 23.1181 and 23.1203; and - adding the phrases “adjacent areas that would be subjected to the effects of fire in the designated fire zones.” ### Page 147 135 The intent of this section in the rules is to require that the materials and components, in the designated fire zone, that are essential to flight safety be fabricated from a material meeting the definition of fireproof under Title 14 Part 1.1 or be shown to be capable of maintaining their integrity or performing their function under the conditions of fire at least as well as steel. The intent of this section is also to require that materials and components, in adjacent areas to a designated fire zone, that are essential to flight safety, be capable of maintaining their integrity or performing their function under the conditions of fire in the designated fire zone. The fire condition characterized by a 2000° F flame can be treated as a failure condition that should not prevent continued safe flight and landing for at least fifteen (15) minutes. The rule requires the structures (composite and metallic) behind the firewall and subjected to the heat effects of the fire be able to withstand the flight loads expected to occur during completion of the flight. It also requires these flight loads to not be less than the gust loads expected to be encountered during the completion of the flight. These loads can be treated as ultimate loads. Design features, including multiple load path arrangement, can be taken into account when establishing the remaining structural capacity. Freedom from flutter and whirlmode should also be demonstrated. Compliance with the above requirement must be demonstrated by tests, or by analysis supported by tests. The assessment of heat effects needs to include all heat transfer mechanisms that may occur in the area of concern. For composite structure, the longterm environmental effects that may degrade the mechanical properties of the structures also need to be considered. These may include the effects due to moisture and steam pressure. JAA ACJ 23.865 is acceptable for FAA certification. ### Page 148 **ELECTRICAL BONDING AND LIGHTNING PROTECTION** 23.867 Electrical bonding and protection against lightning and static electricity Amendment 23-7 and Subsequent Lightning protection of VFR airplanes was considered because there is a possibility that a lightning strike on a VFR airplane could occur. However, the probability and consequences of a VFR lightning strike are more pertinent than the possibility. The hundreds of millions of hours of service history on metallic airplanes illustrate neither a probability nor a consequence worthy of requiring the customer’s assets be expended on lightning certification of this class of airplane. Therefore, this section is not applicable to VFR-Only airplanes that have electrical-bonding characteristics commensurate with metallic construction. Protection of Composite Structure and Installed Equipment from the Direct and Indirect Effects of Lightning Unless VFR-Only, per the preceding paragraph, an applicant should submit a certification test plan describing the analyses/testing to be used to demonstrate lightning protection effectiveness. At a minimum, the test plan should describe the sub-components, components, and systems to be tested and include the following: a. Zone definition of the entire aircraft (show how this will be accomplished—strike attachment model tests, similar design, etc.). b. Full scale or other acceptable means of simulated lightning qualification test on all flight critical portions of the airframe, flight controls, fuel, propulsion, electrical and avionics systems, including damage assessment procedures. c. The indirect effect on the airplane and its systems after a direct lightning strike must also be evaluated. This will include both upset and damage assessment to the electrical and avionics systems. d. Values for each phase of the protection should be delineated. This includes test procedures, test parameters, methodology and simulation techniques to be used during the validation phase. e. Grounding and bonding concerns are paramount in any aircraft but are more critical in composite structures. The test plan must address these areas and include life cycle environmental tests. ### Page 149 f. Lightning simulation methods including test voltages and current waveforms must be defined. g. The test plan should address maintenance practices and repairs to the structure that ensure the continued lightning protection effectiveness. The following SAE documents are applicable: SAE ARP 5412, Aircraft Lightning Environment and Related Test Waveforms. SAE ARP 5413, Certification of Aircraft Electrical/Electronic Systems for the Indirect Effects of Lightning, (being converted to an AC, AC 20-136A). SAE ARP 5414, Aircraft Lightning Zoning. SAE ARP 5415, User’s Manual for Certification of Aircraft Electrical/Electronic Systems for the Indirect Effects of Lightning. ### Page 150 **MISCELLANEOUS** 23.871 Leveling means No policy available as of June 30, 2001. ### Page 151 139 Subpart F—Equipment **GENERAL** 23.1301 Function and installation Original Issue and Subsequent When performing Technical Standard Order (TSO) testing to RTCA/DO-160, Sections 16 and 17, the sentence “After exposure. DETERMINE COMPLIANCE WITH APPLICABLE EQUIPMENT PERFORMANCE STANDARDS” is considered to refer to the appropriate section of the RTCA Minimum Operational Performance Standards (MOPS) that references DO-160. For a test section of a MOPS that is titled “Normal Operating Conditions,” the applicable equipment standards should be met while the test is being conducted. If the equipment experiences an interruption of operation as a result of the test, then the acceptability of the duration of the interruption will depend on the intended function of the equipment and must be justified. Manual resets or pilot actions required to restore normal operation following any test must be explicitly permitted by the applicable TSO or MOPS. If the TSO or referenced MOPS does not specifically allow for a reset, then the equipment must continue to operate or resume normal operation without any pilot action. Instruments and equipment installed to meet the requirements of 14 CFR part 23, § 23.1303 and part 91, § 91.205 should meet the Technical Standard Order requirements that are applicable for that instrument or equipment or equivalent requirements. Original Issue through Amendment 23-19 A system/equipment that is neither essential for safe operation nor required by airworthiness or operating rules may be approved if it is not a hazard in normal operation or when it malfunctions/fails. It does not have to perform its intended function. Section 23.1301 requires that instruments be installed in accordance with prescribed limitations. Therefore, if an instrument manufacturer specifies any allowable installation requirements (i.e., panel slope for gyroscopic instruments), the installer should stay within the limitation(s). Each piece of installed equipment must be labeled as to its identification, function or operating limitations or any combination. This applies to the manufacturer of the equipment ### Page 152 140 not the installer. The installer is required to verify the intended function and make any placards or flight manual limitations per Subpart G that the installed equipment makes necessary. See AC’s 23-8A, Flight Test Guide for Certification of Part 23 Airplanes; 20-67B, Airborne VHF Communications Equipment Installations; and 20-41A, Substitute Technical Standard Order (TSO) Aircraft Equipment. Amendment 23-20 and Subsequent All installed systems/equipment should perform their intended functions. For systems/equipment neither essential for safe operation nor required by airworthiness or operating rules, the manufacturer should define the intended functions that the FAA will verify as part of the certification project. Section 23.1301 requires that instruments be installed in accordance with prescribed limitations. Therefore, if an instrument manufacturer specifies any allowable installation requirements (i.e., panel slope for gyroscopic instruments), the installer should stay within the limitation. We recommend that the slope be no more than 15°. If applicants want a slope greater than 15°, they should show conclusively by tests or analyses that the instrument will function properly when subjected to all expected airplane maneuvers. Each piece of installed equipment must be labeled as to its identification, function or operating limitations or any combination. This applies to the manufacturer of the equipment not the installer. The installer is required to verify the intended function and make any placards or flight manual limitations per Subpart G that the installed equipment makes necessary. There has been a trend to install equipment, mainly navigation related such as moving maps, as non-required, “Not approved for primary navigation” or “Situation Awareness Only (SA- Only).” The basis for certification has been to perform its intended function and not present a hazard per this section. Instruments that aid situational awareness should be certified per the requirements in § 23.1309, functional hazard assessment. It is not acceptable to label an instrument as “SA-Only” and assume its failure in normal operation is acceptable. It is also unacceptable to place such an instrument in the primary field of view of the pilot. Based on history where pilots have tended to fly to known failed gyroscopic instruments, they cannot be expected to ignore an instrument in this position. Last, it is not acceptable to install an instrument as non-required and use its outputs as input data to required instruments. Guidance for Required Instruments The purpose of this guidance is to provide clarification on instrument and equipment requirements that are required by 14 CFR part 91, § 91.205 for part 23 airplanes. ### Page 153 141 Instruments and equipment installed to meet the requirements of § 91.205 should meet the Technical Standard Order (TSO) or equivalent requirements. Technical Standard Order or equivalent requirements are an acceptable means of compliance for the instrument or equipment standards for installation in part 23 airplanes. The pertinent requirements in 14 CFR part 23 for the basis of the above statement are as follows: (1): Section 23.1301 Function and installation. Each item of installed equipment must-- (a) Be of a kind and design appropriate to its intended function. (b) Be labeled as to its identification, function, or operating limitations, or any applicable combination of these factors; (c) Be installed according to limitations specified for that equipment; and (d) Function properly when installed. And (2): Section 23.1525 Kinds of operation. The kinds of operation authorized (e.g., VFR, IFR, day or night) and the meteorological conditions (e.g., icing) to which the operation of the airplane is limited, or from which it is prohibited, must be established appropriate to the installed equipment. Accuracy of the Magnetic Gyroscopically Stabilized Heading System PURPOSE. The purpose of this guidance is to provide information on accuracy of the magnetic gyroscopically stabilized heading system that can be displayed on a Horizontal Situation Indicator (HSI) installed in part 23 airplanes. The operating rules such as 14 CFR part 91 and 14 CFR part 135 specify the minimum required equipment that must be installed based in part 23 airplanes for the type of operation, such as Visual Flight Rules (VFR) or Instrument Flight Rules (IFR). 14 CFR part 91, § 91.205 requires for heading information; under VFR operation, a magnetic non-stabilized direction ### Page 154 142 indicator (i.e. compass) is required and; in addition under IFR operation, a gyroscopically stabilized heading system is required. The general airworthiness requirements in 14 CFR part 23, § 23.1301 and § 23.1525, determine the flight instrument and equipment accuracy requirements for part 23 airplanes. Part 23 does not prescribe specific accuracy requirements for magnetic gyroscopically stabilized heading systems. Specific accuracy requirements for avionics may be found in Technical Standard Orders and as acceptable means of compliance to § 23.1301 in advisory circulars, notices, or policy statements/letters. ACCURACY REQUIREMENTS: Magnetic Non-Stabilized Direction Indicator. A magnetic non-stabilized direction indicator (compass) is required (reference § 23.1327) to have an accuracy of ±10º or have a correction card, placard, or a back-up gyroscopic direction indication provided that indicator has an accuracy better than ±10º. If the sole purpose of the gyroscopic direction indication is for backing up the magnetic non-stabilized direction indicator, then the accuracy of displayed headings can also be to ±10º. However, if a gyroscopic direction indicator is installed to meet the IFR operating rules, then the installation requirements are defined by § 23.1301. Magnetic Stabilized Gyroscopically Stabilized Direction Indicator. An installed final accuracy for a magnetic stabilized gyroscopically direction indicator of ±4º on the ground or ±6° in normal level flight on any heading would meet the requirements of 14 CFR part 23, § 23.1301. This accuracy applies after compensation and should include cumulative errors in all combinations due to: • the equipment itself, • the current flow in any item of electrical equipment and its associated wiring, • the movement of any component, (e.g. controls or undercarriage), and • the proximity of any item of equipment containing magnetic material direction indicators COMPARATOR MONITOR For systems installations that include two magnetic gyroscopically stabilized heading systems and a comparator that monitors the differences between the headings of the two systems, the comparator trip point set as follows would meet the requirements of 14 CFR part 23, § 23.1301: ### Page 155 • 6º in stabilized level flight. • 6º plus ½ of the bank angle; or • 12º with a bank angle greater than 6º ; • The alert function can be disabled at a bank angle greater than 20º. • An alert is provided if the condition exceeds 60 seconds, but allow two minutes for a turn error as stated in the TSO. It should be noted that the 6º trip point during level flight actually permits a heading error of as much as 12º. This would be comprised of one system at the 6º in-flight tolerance limit while the other system, presumably with some sort of malfunction, could have an error of 12º in the same direction before the comparator monitor alert is tripped. BACKGROUND INFORMATION. The following background information is provided for additional information. **EQUIPMENT REQUIREMENTS.** Instruments and equipment installed to meet the requirements of § 91.205 should meet the Technical Standard Order (TSO) or equivalent requirements. Technical Standard Order or equivalent requirements are an acceptable means of compliance for the instrument or equipment standards for installation in small airplanes. The pertinent requirements in 14 CFR part 23 for the basis of the above statement are as follows: Section 23.1301, Function and installation. Each item of installed equipment must-- (a) Be of a kind and design appropriate to its intended function. (b) Be labeled as to its identification, function, or operating limitations, or any applicable combination of these factors; (c) Be installed according to limitations specified for that equipment; and (d) Function properly when installed. And Section 23.1525, Kinds of operation. The kinds of operation authorized (e.g., VFR, IFR, day or night) and the meteorological conditions (e.g., icing) to which the operation of the airplane is limited, or from which it is prohibited, must be established appropriate to the installed equipment. **VFR REQUIREMENTS.** A magnetic non-stabilized direction indicator is required equipment by § 23.1303 and also for VFR operation by § 91.205. The required accuracy for magnetic non-stabilized direction ### Page 156 144 indicator is prescribed in § 23.1327 as ±10º unless a correction card, placard, or gyroscopic direction indicator is used. If the sole purpose of the gyroscopic direction indication is for backing up the magnetic non-stabilized direction indicator, then the accuracy of displayed headings can also be to ±10º. This should not be interpreted as the required accuracy for the gyroscopic direction indicator in general. **IFR REQUIREMENTS.** For IFR, a gyroscopic direction indication is required by § 91.205. The gyroscopic direction indicator should meet the TSO-C5d or TSO-C6c or equivalent. A Direction Instrument, Magnetic (Gyroscopically Stabilized) should meet the minimum performance standard of TSO- C6d. As with most TSO's, TSO-C6d refers to SAE documents AS 8013 for the minimum performance standards. In this document, the standard in paragraph 4.4, Turn Error, is “Two minutes after resumption of straight and level flight the scale error resulting from coordinated turn of 180º in 1 minute at a true airspeed of 180 miles per hour (289 km/h) shall be within 2º.” During stabilized and under environmental conditions, the test requirement for the error tolerance is 2º. The TSO is an equipment minimum performance standard and not an installation standard. For the TSO qualification, the equipment is tested under extreme environmental conditions and they should be compatible for this airplane installation. See AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes, for additional guidance. ### Page 157 23.1303 Flight and navigation instruments Original Issue and Subsequent Altimeters A servo-corrected altimeter may be installed as the required altimeter provided an electrical failure is apparent to the pilot and the altimeter meets the accuracy requirements of the standard pneumatic altimeter without electrical power. Or, a servo-driven or servocorrected altimeter with insufficient accuracy may be installed with at least one pneumatic altimeter installed for use by the pilot. On aircraft requiring two pilots, instruments should be located in front of each pilot. Therefore, either the pneumatic or the electrical altimeter can be installed in either location. The desired level of safety could be achieved without a pneumatic altimeter if the electrical supply is ensured. The provision of a pneumatic altimeter is usually more practical than the design and installation of a suitably reliable electrical supply system. Altimeters that employ a “Smiths Law” correction are acceptable provided they are identified by an appropriate part number, marked clearly for use only on the airplane on which they are calibrated, and information is available to the pilot to enable manual correction computations at airspeeds other than those used in designing the instrument correction. Amendment 23-17 and Subsequent See AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes, for information on Free Air Temperature Instruments and Speed Warning Devices required for turbine engine powered airplanes. Altimeters For installation of electronically powered altimeters, when the regulations were promulgated for the requirements of altimeter systems, only pneumatic altimeters were envisioned. The minimum level of safety established by the regulation was based on the reliability and failure modes of pneumatic altimeters. a. Service history has shown numerous occurrences of complete loss of primary electrical power for both single-engine and multiengine airplanes. The complete loss of altimeter information from a failure of primary power could adversely affect the safe operation of the airplane and is considered an unsafe feature. An electrically powered altimeter installation should have a level of safety equivalent to a pneumatic altimeter installation, and it may be found acceptable if there is no unsafe feature or characteristic. ### Page 158 b. In assessing an electrically powered altimeter with pneumatic reversion capability, the means of providing continuous and usable altitude information should be considered upon a failure of the primary electrical power. An electrical powered altimeter may be acceptable under one of the following types of installation: (1) An electrical powered altimeter with pneumatic reversion that provides a power failure warning as an integral part of the instrument’s display, and appropriate correction information is provided for the reversionary pneumatic mode. (2) An electrical powered altimeter that is provided with an alternate power source independent of the electrical generating system. Adequate information should be provided to the pilot on the operating limitations and procedures when operating on the alternate power source. (3) An electrical powered altimeter without a pneumatic reversionary mode may be installed at any pilot’s position provided a pneumatic altimeter is located on the instrument panel so that it is found to be usable from any pilot’s flight position. Guidance Regarding Latent Failures of Altimeters Though “pilot error” seems likely as a primary cause in a controlled flight into terrain (CFIT) accident, we recognize that CFIT accidents are unfortunately common and a misleading, latent failure of an altimeter could impact other airplanes. Rulemaking to require monitoring of an encoding altimeter to activate a warning means for a failure is possible. Whether this is economically justified will have to be determined. The following options would be useful in a single pilot, high workload situation such as a descent in Instrument Meteorological Conditions. The Small Airplane Directorate recommends that applicants do one of the following when installing an altitude encoding altimeter: 1. Install a monitor between another altitude source and the primary, encoded altimeter. If installed, a Global Positioning System (GPS) source would be best. Another altimeter would also be acceptable. A warning would indicate a discrepancy and permit troubleshooting to determine which altitude source is failed, or 2. Install a non-encoding altimeter as primary with the encoding altimeter as secondary. This would allow for warnings from air traffic controllers about terrain clearance when primary system failures were undetected by cross checks, or ### Page 159 3. Install a Terrain Awareness Warning System (TAWS). TAWS Class A & B are required for turbine-powered airplanes configured with six or more passenger seats operating under parts 91 and 135 and all turbine-powered airplanes operating under part 121. (See AC 23-18 for more details). We plan to initiate a rulemaking project for Item 1. LORAN-C See AC 20-121A, Airworthiness Approval of Airborne Loran-C Navigation Systems for Use in the U.S. National Airspace System (NAS), for information on Loran-C installations. Global Positioning System (GPS) See AC 20-138, Airworthiness Approval of Global Positioning System (GPS) Navigation Equipment for Use as a VFR and IFR Supplemental Navigation System. For GPS based navigation systems, the maximum allowable time to reestablish a valid navigation position is five seconds. This reacquisition period is considered normal operation; the navigation failure flag(s) or annunciation(s), therefore, should not be displayed and the equipment must not present misleading information to the flight crew. Questions have arisen regarding the intent of paragraph (a)(3)(x) of TSO-C129 as it is applies to approach qualified (Class A1) GPS receivers. The paragraph states in part that the databases would contain all Standard Instrument Departures (SID's) and all Standard Terminal Arrival Routes (STAR's). The initial intent was that these procedures be selectable by name and not require manual selection of all waypoints. The workload associated with manual selection of individual waypoints has been determined to be unacceptable for single pilot operation. There are two potential deviations to the original intent of the TSO that must be requested formally by the TSOA Holder: - If the SID's are not stored and selectable by name, then a limitation must be placed in the airplane flight manual supplement stating that SID's must be entered prior to departure. - If all named waypoints of the STAR's are stored individually in the database (not selectable as a procedure), then a limitation must be placed in the airplane flight manual supplement stating that GPS equipment is not approved for conducting STAR's. This limitation may be removed as part of an installation approval for aircraft requiring more that one pilot crewmember. ### Page 160 148 GPS as a Primary Means of Navigation See Action Notice N8110.60, GPS as a Primary Means of Navigation for Oceanic/Remote Operations, and the following clarifications: Paragraph 6.d states that follow-on approvals in a different make/model aircraft require an STC. An engineering review is necessary to establish compliance with the requirement that the loss of navigation is demonstrated to be improbable (paragraph 4f of N8110.60). This may require dual GPS antennas. As a supplemental system, GPS may be used as one of two long-range navigators where two are required for oceanic/remote operations and the other is a sole-means navigation system (e.g., INS, Omega). Supplemental GPS may also be used as the only long-range navigation system for oceanic/remote operations that require one long-range and one short-range navigation system (Reference AC 90-94, Section 1, paragraph 3b); these operations, therefore, do not require GPS equipment that complies with N8110.60, and fault detection and exclusion (FDE) is not required. For these operations, GPS equipment should satisfy the requirements described in TSO-C129a and be installed in accordance with AC 20-138 or AC 20-130A. As stated in paragraph 6b of the Notice, no additional tests above those specified in AC 20-138 and AC 20-130A are required. New installations of primary means GPS equipment can be treated like supplemental equipment except that the loss of the longrange navigation function must be demonstrated to be improbable (paragraph 4f of the Notice). Some GPS manufacturers may wish to demonstrate that their previously approved articles comply with N8110.60. They must obtain a letter of design approval as described in paragraph 6a of N8110.60 and send a copy to all their customers who own the compliant equipment. Since the equipment was previously installed and approved and there are no modifications, none of the qualification procedures in AC 20-138 or AC 20-130A are needed. However, the installation will still have to be certified for compliance with the requirement that loss of the long-range navigation function is improbable. The Airplane Flight Manual Supplement (AFMS) should be updated per paragraph 6c. If the second GPS unit was not originally installed and approved, the second unit should be evaluated in accordance with AC 20-138 or AC 20-130A, as applicable. Some GPS manufacturers may develop an upgrade to a TSO Authorized GPS unit. This upgrade is considered a major design change as described in 14 CFR part 21, § 21.611(b). They must obtain a letter of design approval as described in paragraph 6a of N8110.60 and include a copy with each upgraded unit. Provided there are no other changes to the article's design, none of the qualification procedures in AC 20-138 or AC 20-130A are needed. However, the installation must still be certified as the upgrade is a ### Page 161 149 modification to the approved type design. The certification will evaluate compliance with the requirement that the loss of the long-range navigation function is improbable. The AFMS should be updated per paragraph 6c. If the second GPS unit was not originally installed and approved, the second unit should be evaluated in accordance with AC 20- 138 or AC 20-130A, as applicable. Paragraph 4g of N8110.60 requires an FDE prediction capability, and the notice defines an algorithm for calculating the acceptable duration of an exclusion outage. This capability should allow the user to specify the maximum outage duration for the assessment of whether the operation can be conducted, since additional criteria may be specified for specific operations (e.g., RNP operations). Paragraph 5a recommends that the GPS equipment be capable of using GPS satellites down to a mask angle of zero degrees. Installed, achieved mask angle will frequently be higher due to installed antenna gain characteristics, signal blockage, and aircraft pitch during level flight. For these reasons, a mask angle of less than two degrees should not be used for prediction capability. Installations previously approved using a lower mask angle do not need to be re-evaluated. Terrain Awareness Warning System (TAWS) This guidance is in response to a request for clarification on follow-on installation field approval by the Flight Standards District Office (FSDO) for the Terrain Awareness and Warning System (TAWS), Class B, for part 23 airplanes. The questions regarded field approvals by the Airworthiness Aviation Safety Inspector (ASI) approving an Airplane Flight Manual Supplement (AFMS), or Supplemental Airplane Flight (AFM) Manual. We have coordinated this guidance with the Continuous Airworthiness Maintenance Division, Flight Standards Service (AFS-300). Flight Standards Information Bulletin for Airworthiness (FSAW) 02-03A, Follow-On Approval of Class B Terrain Awareness and Warning Systems (TAWS) (Amended) was issued April 16, 2002. This bulletin explains the standards a Class B Terrain Awareness and Warning System (TAWS), also called an Enhanced Ground Proximity Warning System (EGPWS), must meet to qualify for a follow-on field approval. Policy for TAWS B Displays of Geometric Altitude Labeled Mean Sea Level (MSL) The Small Airplane Directorate was recently informed that some Aircraft Certification Offices (ACO's) have stopped certification projects of Terrain Awareness Warning Systems (TAWS) with electronic displays. The Small Airplane Directorate did not issue any policy, etc., to stop any installations of TAWS B in part 23 airplanes. This policy is intended to permit current and future installations of TAWS equipment, which is not only required for some part 23 airplanes, but also proven to enhance safety. ### Page 162 150 The Small Airplane Directorate’s position is that the TAWS B system with a display is a safety improvement over a warning only system in reducing CFIT accidents/incidents. The display provides two significant enhancements: 1) improved situational awareness, and 2) fewer altimeter errors (e.g., due to ground-pilot communication/interpretation errors, pilot setting errors, static source errors, static system failures, icing effects, pressure and temperature changes, non-standard pressure gradients, etc.) However, as the current universal standard, pilots must use their primary barometric based altimeter instrument for navigation within the National Airspace System. Thus, the MSL label on the TAWS altitude display, without clarification of its geometrically calculated multisensor source, could be misleading due to customary pilot expectations of a barometric based altimeter source. Nonetheless, at this time there is no history of accidents or incidents from use of the MSL label on this geometric altitude. Because this label has not been shown to be a safety hazard, the FAA cannot mandate a corrective Airworthiness Directive for previous installations. Therefore, because the need or nature of a future labeling change is still being debated within the FAA, and until such time that a change may be required, current and future installations in part 23 aircraft may continue to be approved. These installations will require agreed-to additions to airplane flight manuals and TAWS user guides as per the following paragraphs. 1. The Airplane Flight Manual (AFM) or AFM Supplement (AFMS) will describe the following limitations of the TAWS altitude displayed: The indication of MSL altitude on the upper left-hand corner of the Terrain Awareness Display must not be used for navigation, especially for maintaining an ATC assigned altitude. Navigation must not be predicated upon the use of the Terrain Awareness Display. The Terrain Awareness Display is intended to serve as a situational awareness tool only and may not provide the accuracy and/or fidelity on which to solely base terrain or obstacle avoidance maneuvering decisions. 2. The Airplane Flight Manual (AFM) or AFM Supplement (AFMS) will describe the following Normal Procedures of the TAWS altitude displayed. “The indication of MSL altitude is shown on the upper left hand comer of the Terrain Display. This altitude is the reference altitude for the display and the terrain awareness algorithm. This reference altitude is based on internally calculated Geometric Altitude and NOT corrected barometric altitude that must be used when navigating within the National Airspace System. Geometric Altitude is the height above mean sea level (MSL) derived from the GPS receiver, filtered by the vertical figure of merits from the ### Page 163 151 same GPS and complemented by short term variations in barometric altitude. It represents the aircraft's calculated true height above MSL and serves as the reference altitude for color-coding of the terrain display and the altitude input to the look-ahead algorithm. Because it is primarily comprised of GPS altitude, this reference altitude will often differ from cockpit displayed corrected barometric altitude. The geometric altitude is not to be used for navigation. It is presented to provide the crew with additional situational awareness of true height above sea level upon which terrain alerting and display is based. GPS altitude is an altitude above mean-sea-level and it is the geodetic height above the WGS-84 ellipsoid corrected by the geoid height in the GPS receiver itself. With Selective Availability turned off as currently, the accuracy is usually better than 75 feet and with Selective Availability turned on, short term accuracy is in the order of 400 feet, but the geometric altitude should be within 100 feet.” 3. The TAWS manufacturers will provide the above information to all installations until or unless they develop a software change that will re-label geometric altitude something else such as “GPSA.” ### Page 164 23.1305 Powerplant instruments Original Issue and Subsequent a. Fuel Pressure Indication Paragraph (b)(4)(ii) of § 23.1305 requires fuel pressure indicators for pump-fed engines. An equivalent level of safety finding can be made for a warning (red per § 23.1322) light set to operate when the primary pump fails and the emergency pump must be manually activated. A caution (amber per § 23.1322) light is acceptable for an automatic switchover to the emergency fuel pressure pump. Also, a fuel flow indicator can be used to indicate the primary pump is operating normally if there is a placard or Airplane Flight Manual (AFM) to advise the pilot on how to determine primary pump condition from fuel flow information. b. Powerplant Instrument Marking See AC 20-88A, Guidelines on the Marking of Aircraft Powerplant Instruments (Displays). In consideration of the policy in Item 6d of AC 20-88A, where the rate of change is small or nearly steady state (i.e., cylinder head temperature, exhaust gas temperature, or turbocharger inlet temperature), reciprocating engine parameter instruments may use direct reading digital (alphanumeric) instrument displays with ancillary displays such as warning lights. These ancillary light displays should include amber lights for takeoff/cautionary ranges and red lights for appropriate limits. Placards containing operating range and limitation information should also be included. c. Fuel Flowmeters This guidance is applicable to the installation of fuel flowmeters in small airplanes with continuous-flow, fuel injection, and reciprocating engines. 1. RELATED REGULATIONS These acceptable means of compliance refer to certain provisions of part 23 and the corresponding provisions of the former part 3 of the CAR in the case of airplanes for which those regulations are applicable. Listed below are the applicable part 23 sections with the related CAR sections shown in parentheses: Part 23 Sections § 23.773 (3.382) § 23.955 (3.433) § 23.961 (3.438) ### Page 165 153 § 23.991 (3.449) § 23.993 (3.550) § 23.1183 (3.638) § 23.1191 (3.624) § 23.1305 (3.655) § 23.1337 (3.673) § 23.1529 § 23.1541 (3.755) § 23.1543 (3.756) § 23.1549 (3.759) **2. BACKGROUND** a. Recently there has been a trend toward replacing fuel pressure indicators and analog reading fuel flowmeters with digital fuel flowmeters/fuel totalizers. New developments in microprocessor technology have resulted in digital fuel flow computer systems that are economical, accurate, and that provide data for improved fuel management. These digital fuel flow computer systems also have features for displaying total fuel consumed, total fuel remaining, and time remaining; however, the accuracy of these readings is dependent upon the initial fuel supply entered into the fuel computer. The precise digital readings that are displayed to the nearest tenth of a gallon could give a pilot a false sense of accuracy and security, especially the readings for total fuel remaining and time remaining. b. Digital fuel flowmeters are not a required powerplant instrument except for turbine engine airplanes with an Amdt. 23-43 certification basis. They are optional equipment and should not be considered replacements for fuel quantity or fuel pressure indicators. Different interpretations of the regulations have caused conflict and lack of national standardization on installation of fuel pressure indicators and fuel flowmeters/fuel totalizers in small airplanes that have continuous-flow, fuel-injection systems in reciprocating engines. Inquiries from members of the aviation community and manufacturers have indicated a need for information concerning approval and installation of digital fuel flowmeters/fuel totalizers. The location of the fuel flow transducer in the fuel system is critical for measuring the total fuel flow consumed by the engine and maintaining engine performance. Each type of installation has an impact on the operation of the fuel system and needs to be evaluated and approved. **3. DISCUSSION** a. Fuel Pressure and Fuel Quantity Indicator (1) A fuel pressure indicator is required for pump-fed engines in accordance with § 23.1305(g). It is intended to monitor metered fuel pressure at the inlet to the ### Page 166 154 injector and to advise the pilot of a fuel pressure deficiency. Many small airplanes with reciprocating, continuous-flow, fuel-injection engines are equipped with fuel pressure indicators that actually measure metered fuel pressure. Metered fuel pressure in a fuel-injection system also relates to fuel flow, and can provide a satisfactory method for displaying fuel flow. However, replacing the metered fuel pressure indicators with fuel flowmeters could cause an unsafe condition by failing to provide critical fuel pressure information to the pilot that is especially important during the takeoff phase of flight. Fuel flowmeters are not required powerplant instruments for reciprocating engines to meet airworthiness standards of part 3 of the CAR or part 23. (2) Digital fuel flow computer systems have a fuel flow transducer that directly measures the amount of fuel being fed to the engine. The fuel flow transducer may be a small paddle wheel, an impeller, or spring-loaded movable vanes. Digital displays with a fuel computer also permit these instruments to display total fuel consumed, total fuel remaining, and time remaining at the present fuel flow rate for fuel management. Overall accuracy for fuel remaining and time remaining readings depends on the transducer processing unit and display. The largest possible error is the initial fuel supply, which is entered by the pilot at the start of each flight. Errors in the initial fuel supply may be caused by an uneven ramp, unusual loading, volume changes of the fuel due to temperature variations, malfunctions in the fuel system such as leaks, siphoning actions, collapsed bladders, and other factors. Consequently, total fuel remaining should be verified with the fuel quantity indicator. In accordance with § 23.1337(b)(1), fuel quantity indicators are required to be calibrated to read "zero" during level flight when the quantity of fuel remaining in the tank is equal to the unusable fuel supply. For this reason, fuel quantity indicators should be used as the primary fuel-remaining instruments. Fuel quantity indicators that are inaccurate should be periodically calibrated, repaired, or replaced, as necessary, to ensure reliable readings. b. Fuel-Injection Systems Fuel-injection systems have been designed for many types of reciprocating engines, and they vary in details of construction, arrangement, and operation. Only continuous-flow, fuel-injection systems for reciprocating engines will be discussed in either the speed-sensing pressure pump or constant-pressure pump categories. (1) Fuel-Injection System with Integral Speed-Sensing Pressure Pump (a) A fuel-injection system with an integral speed-sensing pressure pump delivers fuel at a pressure proportional to engine speed, and the pump is approved as part of the engine type design during the engine certification process. The fuel-injection system has fuel lift capability that enables the system to function ### Page 167 155 with a negative inlet pressure within specific limits as indicated by the engine type data sheet. An emergency fuel pump is not required when the fuel injection pump is approved as part of the engine in accordance with § 23.991(b). The airframe manufacturers may provide an auxiliary fuel pump located upstream of the fuel-injector pump for priming the engine and suppressing fuel vapors. This auxiliary fuel pump can provide some fuel during emergency operations but may not sustain engine operation at full power in the event the engine-driven, fuel-injector pump fails; therefore, it is not considered an emergency fuel pump. (b) If the fuel system in the airplane can meet the fuel flow requirements of § 23.955(c) at the minimum allowable inlet pressure limits without the need of an external pump, a fuel pressure indicator is not required. Nonetheless, some manufacturers have installed a fuel pressure indicator that senses metered fuel pressure at the fuel distribution valve. Since metered fuel pressure is related to fuel flow, it can provide a means for displaying fuel flow. A pressure indicator that is measuring metered fuel pressure may have the scale marked in terms of fuel pressure, fuel flow, or percentage of engine power. With these fuel flow markings, the indicator sometimes is referred to as an analog pressure-type flowmeter. If an analog pressure-type flowmeter is installed as part of the airplane manufacturer's type certificate, a replacement digital or analog fuel flowmeter/fuel totalizer is acceptable, provided the installation meets the applicable airworthiness requirements mentioned in the Acceptable Means of Compliance. (2) Fuel-Injection System with Constant Pressure Pump (a) A fuel-injection system with constant discharge pressure during normal flightengine-revolutions usually requires that fuel be supplied at a positive pressure within specified limits to the fuel-injector inlet. To provide this inlet pressure, the engine-driven fuel pump and the emergency pump are usually installed by the airplane manufacturer. An emergency fuel pump is required by § 23.991(b), and this pump should meet the fuel flow rate of § 23.955; therefore, it will sustain engine operation if the engine-driven fuel pump fails. (b) A fuel pressure indicator is required for pump-fed engines in accordance with § 23.1305(g) and is intended for monitoring unmetered fuel pressure at the inlet to the injector. The fuel pressure indicator provides a means for the pilot to determine if the fuel pressure is within safe limits for proper operation. (c) Several airplanes have been approved with a fuel pressure indicator connected to the fuel distribution valve where the fuel flow is a function of metered fuel pressure to the discharge nozzle. Metered fuel pressure is related to fuel flow ### Page 168 156 and also relates to engine power output. In some applications, metered fuel pressure has been found acceptable for monitoring fuel pressure and controlling engine performance. The scale on the pressure indicator is to be marked in fuel pressure; in addition, it may be marked in either fuel flow or percentage of engine power output. A fuel pressure indicator at the inlet to the injector provides a more positive means of monitoring the operation of the engine-driven fuel pump and the emergency fuel pump. (d) An airplane that has both a separate unmetered fuel pressure indicator and an analog pressure-type flowmeter may have the analog pressure-type fuel flowmeter replaced with a digital fuel flowmeter/fuel totalizer. If only an analog pressure-type fuel flowmeter is installed that actually operates from metered fuel pressure, the analog pressure-type fuel flowmeter may not be replaced with a digital fuel flowmeter/fuel totalizer unless another fuel pressure indicator is installed to sense the fuel pressure at the fuel-injector inlet. Or, the analog pressure-type fuel flowmeter may be replaced with a digital fuel flowmeter/fuel totalizer if an equivalent level of safety for the airplane shows that replacing the fuel pressure indicator with a flowmeter will still meet the applicable airworthiness requirement. A finding of equivalent level of safety should substantiate that the instrumentation provided by the fuel flowmeter is satisfactory, reliable, and safe under all reasonably foreseeable operating conditions. 4. ACCEPTABLE MEANS OF COMPLIANCE An acceptable method of compliance with the airworthiness standards for installation of fuel flowmeters in small airplanes with continuous-flow, fuel-injection system, reciprocating engine is described below. a. FAA Approval of Technical Data/Installation. Installation of the fuel flowmeter/fuel totalizer may be approved through Type Certification (TC) or Supplemental Type Certification (STC) for either the airframe or the engine. FAA approval is obtained after the applicant shows that the fuel flowmeter/fuel totalizer will perform its intended functions and ensures that no unsafe features are incorporated. The need for certification approval for the engine will be determined for each particular installation. Certification approval for the engine is not required when the applicant provides FAA approved data that shows an alternate configuration that permits a digital flowmeter with specific instructions. An improper installation not only will jeopardize the safety of the present designs, but could also increase the probability of system failure. Installations should comply with the airworthiness regulations and with the manufacturer's installation criteria. ### Page 169 b. Airworthiness Considerations (1) Fuel-Injection System with Integral Speed Sensing Pressure Pump Installation of a digital or analog fuel flowmeter may replace the analog pressuretype flowmeter. (2) Fuel-Injection System with a Constant Pressure Pump Installation of a digital or analog fuel flowmeter may replace the analog pressuretype flowmeter, provided an unmetered fuel pressure indicator is installed or it has been determined that replacing the fuel pressure indicator with a fuel flowmeter constitutes an equivalent level of safety. (3) General Considerations Changes to the fuel systems should be evaluated for fuel flow rates, maximum allowable pressure drop, hot weather operations, vibration and loads on lines and fittings, fire protection, and powerplant instruments, including effects of glare and reflections on instruments in the pilot compartment. An engineering analysis should be made to ensure good engineering practices are incorporated in the design and that the installation is in accordance with airworthiness standards of the following §§ 23.773, 23.955, 23.961, 23.993, 23.1183, 23.1191, 23.1337 and 23.1529 of part 23. The fire-resistant capability of fuel system components in the engine compartment should be evaluated. The extent and nature of ground and flight evaluations depend upon each particular installation. c. Evaluation Modification of the approved fuel system may have major effects; therefore, an evaluation should be conducted to substantiate continued compliance of the fuel system with airworthiness requirements. FAA approval is issued when all airworthiness requirements are met. The following items should be considered: (1) Fuel flow transducer should measure the total fuel flow under all operating conditions with either the engine-driven or the emergency fuel pumps. Some fuel systems provide an alternate fuel flow path under different operating conditions; for this reason, the fuel flow transducer should be installed upstream of the alternate fuel flow path. (2) Fuel flow transducer should be installed downstream of any bypasses or vent returns to the fuel system. ### Page 170 (3) Maximum fuel pressure drop across the fuel flow transducer (normal and blocked conditions) should be within manufacturer's specifications and airworthiness requirements. Fuel pressure drop may affect the minimum fuel injector inlet pressure. The minimum fuel injector fuel inlet pressure may require redefinition, and the instrument range markings on the fuel pressure indicator may need to be revised. An engine-driven pump and emergency or boost pump may require adjustment to a higher pressure to account for the added restriction of the transducer. The pumps should be tested to ascertain their capability to supply the required fuel flow rate at the higher pressure. Flight tests for turbocharged engines may be required to determine that the minimum fuel injector inlet pressure meets the engine type certificate data sheet at the maximum approved altitude. d. Markings and Placards for Powerplant Instruments AC 20-88A provides guidelines on markings of airplane powerplant instruments. Sections 23.1541, 23.1543 and 23.1549 of part 23 provide the airworthiness requirements for instrument markings and placards. Either the required range marking or placards, or both, should be furnished with the safe operating limits. A placard should be located near the fuel flowmeter/fuel totalizer display with the following statement: “Original equipment fuel quantity indicator is the primary reading of fuel on board the airplane.” e. Airplane Flight Manual (AFM) A flight manual supplement or supplemental AFM or placards, if appropriate, should be prepared by the applicant. The information should be presented for FAA approval in the following sections: (1) Limitation section should include placard information and instrument markings. (2) Normal procedure section should include information on the operation and function of the equipment. Included in this section should be information that the fuel totalizer does not sense the quantity of fuel in the tank and it should not be used as a fuel quantity indicator. The accuracy of total fuel remaining displayed on the fuel flowmeter/fuel totalizer is dependent upon the initial fuel supply programmed into the computer before the start of each flight. Uncertainties about initial fuel supply and total fuel remaining can be due to an uneven ramp, unusual loading, volume changes of the fuel due to temperature variations, malfunctions such as leaks, siphoning action, collapsed bladder, and other factors; therefore, the total fuel remaining should be verified with the fuel quantity indicator. Before flight, it is essential that the pilot determine that the fuel programmed into the computer is the same as the usable fuel on board the airplane. ### Page 171 (3) The emergency procedure section should include any system malfunction that may occur due to electrical power failure and the procedures for verifying proper operation after power outages. (4) If the certification basis does not require an AFM with the airplane, the applicant may provide a supplemental AFM or provide the necessary information to the pilot by means of placards. Amendment 23-7 and Subsequent 1. Digital (Alphanumeric) Instruments See AC 23.1311-1A for guidance on this topic. 2. Torque Meter Markings Markings on torque meters should be as follows: a. The maximum safe operating torque should be indicated by a red radial. b. The green arc should extend across the complete normal operating range. c. Takeoff torque can be indicated by the word “Takeoff” or the letters “T.O.” arranged as a radial with an explanation of their significance in the AFM. 3. Warning Means Instead of Indicators Warning means for § 23.1305: oil quantity measuring device, powerplant ice protection indicating means, fuel system anti-ice indicating means, thrust reverser indicating means, and propeller blade angle indicating means, can be acceptable as an equivalent level of safety. 4. Fuel Strainer or Filter Indicators Acceptable means of compliance for fuel strainers or filter indicators for turbine-engine airplanes are as follows: a. A fuel filter approved under 14 CFR part 33, § 33.67, Amendment 33-6, installed within the engine upstream of the high-pressure engine-driven positive displacement pump or the fuel metering device will comply with the provisions of § 23.997 without an airframe supplied filter. The fuel filter should be capable of sustained operation while operating with water in the fuel as specified in §§ 23.991(c) or 33.67(b)(4). An engine-driven, low-pressure fuel pump may be installed upstream of the fuel filter. If ### Page 172 160 an airframe-mounted filter is not installed, care should be taken to ensure there are no undrainable low spots between the fuel tank outlet and the inlet to the engine. b. A fuel strainer approved under § 33.67, Amendment 33-6, would not require an indicator in the cockpit to indicate the occurrence of contamination before it reaches the capacity of the fuel strainer, as required by § 23.1305(c)(8). However, an indicator on the engine should be installed such that it can be readily inspected for operation prior to flight. Instructions for this inspection should be included in the Preflight Check Procedures in the AFM. c. Turbine engine installations that do not have a fuel filter per § 33.67 should have an airframe mounted fuel strainer to comply with § 23.997. Also, an indicator for contamination before it reaches the capacity of the fuel strainer, as required by § 23.1305(c)(8), should be provided. A pop-out button on the filter is not acceptable for compliance to § 23.1305(c)(8). For reciprocating engines, the fuel strainer should comply with all the requirements of § 23.997. See AC 23.1311-1A for guidance on electronic displays of propulsion parameters. ### Page 173 23.1307 Miscellaneous equipment No policy available as of June 30, 2001. ### Page 174 23.1309 Equipment, systems, and installations Amendment 23-14 and Subsequent The FAA has reviewed the part 1 definition of the word “instrument,” and other data and has concluded as follows: a. Where a light is sufficient, the instrument requirement should be changed to a warning means. b. Where trend information is needed, the word “indicator” should be retained. c. Where point information or steps in a sequence need to be shown, the words should be changed to “indicating means” (i.e., the functioning of the ice protection system). Software Notice N8110.89 provided Guidelines for the Approval of Software Changes in Legacy Systems Using RTCA DO-178B. Theoretically, RTCA/DO-178A and RTCA/DO-178B are not very different in terms of objectives. DO-178A states software life-cycle objectives implicitly, while DO-178B states them explicitly. The applicant should determine which DO- 178B objectives are currently not met. Considering this data, the FAA should determine if the missing objectives would affect airworthiness or safety and may require those objectives to be met. Changes that affect airworthiness per 14 CFR part 21, § 21.93 should be considered. The original System Safety Assessment determined the required design assurances for each aircraft function. Software changes to these functions should be further analyzed to determine if the design assurance of those functions is maintained after the software changes are made. Another approach would be to use service experience as defined in DO-178B, Section 12.3.5, when the software is not being changed for the application. Service experience can be used for commercial-off-the-shelf (COTS) software provided the COTS is not embedded within the application. Unless the applicant has complied with the provisions of DO-178B, Section 12.1.4, the accomplishment summary and any box markings shall not indicate approval of the software to DO-178B. Software development and certification for components with multiple functions and interfaces such as electronic engine controls requires significant effort. Even though the software ### Page 175 163 functional requirements are included in DO-178, the verification of software functions as installed in the airplane is necessary to establish that the software and associated functions are correct and certifiable. See AC 23.1309-1C, Equipment, Systems, and Installations in Part 23 Airplanes, for additional guidance. Airplane Parachute Recovery Systems Use of a parachute recovery system is not covered by part 23. Applications for the installation of this system will require special conditions. Amendment 23-41 and Subsequent Lightning Protection See AC 23.1309-1C, AC 20-136, Protection of Aircraft Electrical/Electronic Systems Against the Indirect Effects of Lightning, and RTCA DO-160D, Environmental Test Conditions and Test Procedures for Airborne Equipment, Section 22, for guidance on lightning certification of IFR airplanes. As part of the ongoing review of natural lightning by the SAE Lightning Committee AE2 and EUROCAE WG-31, the multiple stroke and burst environmental criteria was revised from the AC 20-136 requirement as defined in SAE ARP 5412, Aircraft Lightning Environment and Related Test Waveforms. Several SAE documents have been issued and the FAA finds the following SAE documents acceptable: SAE ARP 5412, Aircraft Lightning Environment and Related Test Waveforms. SAE ARP 5413, Certification of Aircraft Electrical/Electronic Systems for the Indirect Effects of Lightning, (being converted to an AC, AC 20-136A). SAE ARP 5414, Aircraft Lightning Zoning. SAE ARP 5415, User’s Manual for Certification of Aircraft Electrical/Electronic Systems for the Indirect Effects of Lightning. High Intensity Radiated Fields (HIRF) Special conditions will still be required for critical systems for High Intensity Radiated Fields (HIRF), since the words “radio frequency energy” in this rule are not intended to include HIRF. RTCA DO-160D, Section 20, is applicable for bench level testing for HIRF. Electronic Engine Control (EEC) Systems Section 23.1309 does not apply to powerplant systems provided as part of the certificated engines per paragraph (f). This rule did not envision digital engine controls. Special ### Page 176 164 conditions will be required for electronic engine controls without a hydromechanical back-up system that has the reliability and performance of a traditional recip engine without electronic engine controls. The special conditions will require compliance to HIRF and § 23.1309(a) through (e), per the latest amendment, for these digital engine controls. Appropriate engine certification data may be used for airplane certification. Even though an engine control may be certificated as part of the engine, the installation aspects of the installation require certification to part 23. Experience has shown that changes required for aircraft certification may require changes to the engine software with the associated engine certification of those changes. Acceptance of engine mounted and engine certificated FADEC or EEC's does not mean that approval at the airplane level is automatic. Flight-testing is still required to assure all part 23 requirements are met. Software changes initiated by the engine manufacturer (or FADEC or EEC manufacturer) should be validated on each software change as to their effect on part 23 certification requirements. Any and all engine control software changes must be coordinated with the appropriate ACO to assess the impact on part 23 certification requirements. ### Page 177 165 INSTRUMENTS: INSTALLATION 23.1311 Electronic display instrument systems Amendment 23-41 through Amendment 23-48 Attitude and Heading For part 23, there is not a specific requirement that instruments at each pilot station be independent unless two pilots are required by the airworthiness or operational rules. Therefore, both electronic attitude and heading instruments can utilize the same attitude and heading reference source. For a single Attitude Heading Reference System (AHRS), the Airplane Flight Manual should include equipment operating limitations to alert the pilot(s) that a failure of the AHRS could simultaneously affect both attitude and heading instruments. However, a single AHRS may not be acceptable if its affect on an autopilot system is a possible catastrophic failure caused by the AHRS, such as an unannunciated slowover (softover) failure. Amendment 23-41 and Subsequent See AC 23.1311-1A, Installation of Electronic Displays in Part 23 Airplanes, for further guidance. Paragraph (e) in § 23.1311 has definitions specific to this rule for “instruments” and “primary.” The definition of “indicators” is not in the regulations. In a search for the word “indicator,” the text does not disclose a precise definition or the intended use of the word. For § 23.1311, an indicator is a means for displaying information on a parameter (and is not the same as an instrument). The preambles for § 23.1311 were intended to allow advanced and new electronic display technologies on part 23 airplanes and to permit the collection of indicators on a single electronic display. More than one indicator could be depicted on one display. For example, a primary flight display may have indicators for attitude, altitude, airspeed, heading, and navigation. Altitude Digital-Only (alphanumeric) displays for barometric altitude should not be approved. ### Page 178 23.1321 Arrangement and visibility Original Issue and Subsequent When applying this rule to powerplant instruments in multiengine airplanes, ensure there is no confusion as to the engine/instrument relationship. For instance, powerplant instruments for the right engine in a twin-engine airplane may not be placed over, under, or to the left of the left-engine instruments. Amendment 23-14 and Subsequent For all installations, the evaluation should consider the different environmental conditions under which the airplane may be operated as defined by § 23.1559. Basic “T” This rule applied the Basic “T” to standardize flight instrument locations. This was not intended to require a “perfect T.” Also, for all installations, the FAA has always intended that § 23.1321(d) apply to each pilot’s station for both type certification and for any operations for which the airplane is approved. Therefore, when an airplane is type certificated with the “basic T” instrumentation at only one pilot’s station, that airplane is limited to operations where only one pilot is required in accordance with §§ 23.1525 and 23.1583(h). ### Page 179 23.1322 Warning, caution, and advisory lights Policy is available in AC 23.1311-1A. ### Page 180 23.1323 Airspeed indicating system Original Issue and Subsequent Part 135, Operating Requirements: Commuter and On-Demand Operations, requires that IFR airplanes have a heated pitot tube for each airspeed system. In consideration of a fourpound bird strike, the minimum distance between pitot tubes that can be accepted is 14 inches, measured in a straight line. Amendment 23-34 and Subsequent See AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes, for additional guidance. ### Page 181 23.1325 Static pressure system Amendment 23-1 and Subsequent Both VFR and IFR airplanes should meet the requirements of § 23.1325 in paragraph (b)(3) of this regulation because static vent icing can occur during both VFR and IFR conditions with hazardous consequences. The rule provides for either an anti-icing means or an alternate static source. a. If installed, the alternate static source is not restricted to emergency conditions but may be used to monitor the primary static system. b. We suggest marking the secondary static source with the word “Alternate.” c. This rule also requires a correction card in the cockpit if the altimeter changes by more than 50 feet on the alternate source. The correction card does not need to be in clear view of the pilot as long as it is available to a pilot seated in the flight position. An acceptable means for alternate-static-source correction data is located in the Performance Section of Airplane Flight Manuals (AFM) with other airspeed and altimeter calibration graphs. A placard that the correction data is available in the Performance Section, which is noted in the Limitations Section of the AFM, may be provided for additional clarification. The alternate static source is separate, and its correction card should provide correction data for the alternate source only. d. The alternate static source is subject to all parts of § 23.1325, as is the primary static source. See AC 20-124, Water Ingestion Testing for Turbine Powered Airplanes, for guidance on testing the airspeed and static systems for water ingestion susceptibility. Amendment 23-34 and Subsequent See AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes, for additional guidance. ### Page 182 23.1326 Pitot heat indication systems Amendment 23-49 and Subsequent This rule requires a caution annunciation whenever the pitot heat is Off or there is a failed heating circuit in the pitot tube heater. The second annunciation cause is fully justified in that it represents a failure condition. The first can have a positive safety effect if it causes pilots to activate the pitot heat in all environmental conditions. This eliminates the loss of the pitot static system due to the pilot error of failing to operate pitot heat when conditions warrant it. A caution annunciation when the pitot heat is Off has two negative issues: (1) It violates the “dark cockpit” where caution and warning lights only represent failure conditions; and (2) Adherence to a “dark cockpit” will cause pitot heat operation in all environmental conditions, which will shorten the life of the system. The Small Airplane Directorate is proposing rulemaking to delete § 23.1326(b)(1) (the requirement for a caution annunciation when the pitot heat is Off). An aircraft design that does not include a caution annunciation when the pitot heat is Off may be eligible for an Equivalent Level of Safety finding that preserves a “dark cockpit” provided a placard or flight manual prescribes when to operate the pitot heat. ### Page 183 23.1327 Magnetic direction indicator Original Issue and Subsequent If the magnetic compass that is required by § 23.1303 is the only heading instrument, then it should meet the requirements of this section. With an approved secondary system such as a directional gyro, and with an appropriate placard to dictate which electrical devices should be switched off when reading the magnetic compass, per § 23.1547, equivalent safety pursuant to part 21, § 21.21(b)(1) may be shown. Regarding magnetic direction indicators: heading information is considered an essential flight instrument function because its loss could result in reduced capability of the flight crew to cope with adverse operating conditions, especially for IFR flights. The indicator specified in this rule was intended to be a magnetic compass (non-stabilized). The requirement for a magnetic direction indicator existed before remote indicating compasses were available. If a magnetic stabilized direction indicator is installed as an additional instrument, the magnetic non-stabilized direction indicator (magnetic compass) is still required as the primary source of magnetic direction. a. A magnetic direction indicator with remote magnetic sensor can be approved under § 21.21 of part 21 if it can be substantiated that it provides a level of safety equal to that provided by the magnetic compass required by § 23.1303(c). The reliability of the system should consider the effects of loss of the airplane’s electrical system, the performance of the equipment under environmental conditions that may be encountered by the airplane, the integrity of the interface wiring, and the reliability of the components. b. For a magnetically stabilized direction indicator approved under an equivalent level of safety finding, the system should be powered from a source that is independent of a single electrical generating system. This other source should be installed so that it is operative without manual selection after total failure of a single electrical generating system. Dual independent stabilized indicator installations with split electrical bus systems may also be approved on multiengine airplanes under an equivalent level of safety finding. The airplane’s battery is not considered an acceptable source unless the state of charge of the battery is displayed to the pilot. c. The following installation requirements of §§ 23.1327 and 23.1547 are also directly related to approval of either type of magnetic indicator: (1) The accuracy is not excessively affected by the airplane’s vibration or magnetic fields. ### Page 184 (2) Deviations of more than 10° in level flight are not permissible, unless a magnetic stabilized direction indicator which does not have a deviation in level flight greater than 10° on any heading, or a gyroscopic direction indicator, is installed. If a gyroscopic direction indicator is installed, it is subject to compliance with § 23.1301 in that it must perform its intended function. It must, therefore, meet the accuracy requirements of TSO-C5, which represent the minimum allowable performance for gyroscopic direction indicators. (3) A placard should show the calibration of the instruments in level flight with the engine(s) operating and whether the calibration was made with the radio receivers on or off. (4) If deviations of more than 10° caused by operation of electrical equipment are approved, the placard should state which electrical loads or combination of loads would cause deviations of more than 10°. ### Page 185 23.1329 Automatic pilot system Original Issue and Subsequent A single malfunction may not result in a hardover signal in more than one axis. When the result of any single malfunction is shown not to be hazardous (no hardover signals) (slowover signals are acceptable if they are determined to be easily controllable without requiring exceptional skill or strength), then multiple axes being affected is acceptable providing the following: a. The malfunction evaluations are acceptable even with the maximum drive signal due to the limited rate of change authority of the powered controlling element and flight control surfaces; b. The monitor/limiting device is independent of the automatic pilot element; c. The signal is less than the hardover signal due to the monitor/limiting device; and d. An acceptable fault analysis shows the functional hazard of a combined monitor failure and automatic pilot malfunction is not catastrophic, including the following: (1) The functional hazard of a failure of a lockout device/system to inhibit autopilot engagement until the pre-engagement check is successfully completed is hazardous or less; (2) Pre-engagement check of the monitor system is mandatory with either a manual or automatic activation means, and; (3) Automatic pilot authority is not greater than necessary to satisfactorily control the airplane. Alterations of increased engine horsepower (and either engine horsepower or major changes in exterior cowlings and surfaces, etc.), in part 23 airplanes, should consider the compatibility of the autopilot system with the increased horsepower, since the malfunction and performance tests of the autopilot are conducted with a defined amount of engine power. Generally, an increase in engine horsepower beyond 10 percent may adversely affect the autopilot system malfunctions, performance, controllability, and longitudinal stability characteristics. Therefore, flight testing may be necessary to verify that the original approval of the autopilot system is still valid. a. The results of malfunction testing determine which flight condition is most critical. The effects of autopilot runaways are more pronounced at aft center of gravity (c.g.). ### Page 186 174 Furthermore, the phase of flight with the largest contribution to adverse conditions varies with airplane model. b. Airplane longitudinal stability is a factor in autopilot system malfunctions. Generally, there is an inverse relationship between engine horsepower and longitudinal stability. Although the turbine engine installations replacing reciprocating engines may be flat rated, the turbine is capable of producing increased horsepower at higher temperatures and altitudes, which could reduce longitudinal stability. Therefore, autopilot performance, especially the pitch axis hardover malfunction, should be evaluated for acceptability. This policy is also applicable to power increases on airplanes with reciprocating engines, either engine replacement or engine modifications that add a turbocharger. c. Performance and controllability evaluations should be considered, including the configuration of most forward c.g. and minimum autopilot authority. This configuration is used to demonstrate that the airplane can be safely controlled by the autopilot when the control surface hinge moment is the highest and the autopilot controllability is at its lowest during corresponding longitudinal trim and airspeed changes. To show compliance with part 23, § 23.1329, applicable to autopilot system installations in small airplanes, the following is acceptable. **1. RELATED REGULATIONS AND DOCUMENTS** a. Regulations These acceptable means of compliance refer to certain provisions of part 23. They may be used in showing compliance with the corresponding provisions of the former Civil Air Regulations (CAR) in the case of airplanes to which the CAR regulations are applicable. For convenience, the part 3 section reference is shown in parenthesis following the part 23 section reference: § 23.143 (3.106) Controllability and Maneuverability, General. § 23.253 High speed characteristics. § 23.395 (3.231) Control system loads. § 23.397 (3.212) Limit control forces and torque’s. § 23.689 (3.345) Cable systems. § 23.777 (3.384) Cockpit controls. § 23.779 (3.384) Motion and effect of cockpit controls. § 23.1301 (3.651 and 3.652) Function and installation. § 23.1309 Equipment, systems, and installations. § 23.1321 (3.661 and 3.662) Arrangement and visibility. § 23.1322 Warning, caution, and advisory lights. ### Page 187 175 § 23.1329 (3.667) Automatic pilot system. § 23.1351 (3.681) Electrical Systems and Equipment, General. § 23.1381 (3.696 and 3.697) Instrument lights. § 23.1431 (3.721) Electronic equipment. § 23.1555 (3.762, 3.763, and 3.765) Control markings. § 23.1581 (3.77) Airplane Flight Manual and Approved Manual Material, General. § 23.1583 (3.778) Operating limitations. § 23.1585 (3.779) Operating procedures. b. Advisory Circulars AC 21-16D Radio Technical Commission for Aeronautics (RTCA) Document DO-160D. AC 23.1309-1C Equipment, Systems, and Installations in Part 23 Airplanes. c. Technical Standard Order TSO-C9c Automatic Pilots. d. Industry Documents RTCA/DO-160D Environmental Test Conditions and Test Procedures for Airborne Equipment. RTCA/DO-178B Software Considerations in Airborne Systems and Equipment Certification. 2. BACKGROUND AC 23.1329-1, Automatic Pilot Systems Approval, which set forth an acceptable means for showing compliance with the autopilot installation requirements, was issued December 23, 1965. Although AC 23.1329-1 was inadvertently canceled in 1977, criteria essentially equivalent to that contained therein continued to be used to show compliance with the applicable autopilot installation requirements. The airworthiness regulations prescribe the requirements for autopilot installation approval. The following criteria have been applied and found reasonable and acceptable in previous type certification programs for complying with specific sections related to these approvals: a. Compliance with the regulations necessitated the conversion of the force exerted by one pilot to overpower an engaged autopilot into measurable terms when either an ### Page 188 176 autopilot quick disconnect or interrupt switch was not provided. The values in the table under § 23.143 are maximums. There may be circumstances where a maximum force less than 75 pounds is required for safety. For example, if a pilot is trying to overpower a nose-up malfunction during climb and reduce power at the same time, a maximum safe force may be less than 75 pounds. Consequently, these forces, as measured at the pilot's controls, were equated to the following temporary and prolonged forces: (1) The maximum temporary force to overpower the autopilot has not been allowed to exceed 30 pounds in roll (force applied at the rim of the wheel), 50 pounds in pitch, and 150 pounds in yaw. These forces are applicable only to initially overpowering the autopilot system. (2) The maximum prolonged force to overpower the autopilot should not exceed 5 pounds in roll, 10 pounds in pitch, and 20 pounds in yaw. b. A reasonable period of time has been established for pilot recognition between the time a malfunction is induced into the autopilot system and the beginning of pilot corrective action following hands-off or unrestrained operation. The following time delays have been acceptable: (1) A 3-second delay following pilot recognition of an autopilot system malfunction, through a deviation of the airplane from the intended flight path, abnormal control movements, or by means of a reliable failure warning system in the climb, cruise, and descent flight regimes. (2) A 1-second delay following pilot recognition of an autopilot system malfunction, through a deviation of the airplane from the intended flight path, abnormal control movements, or by means of a reliable warning system, in maneuvering and approach flight regimes. 3. ACCEPTABLE MEANS OF COMPLIANCE The following procedure, in accordance with the forces and times above, is acceptable as a means of showing that an autopilot system installation is in compliance with the airworthiness rules: a. Cockpit Controls Evaluation of cockpit controls should include the following: (1) The location of autopilot system controls should be readily accessible to the pilot, or both pilots, if a minimum of two pilots is required. ### Page 189 (2) Annunciators should conform to the proper color as specified in § 23.1322. (3) A determination that the controls are usable under bright sunlight and night lighting conditions (§ 23.1381). (4) Either a quick disconnect or interrupt switch for the autopilot system are located on the side of the control wheel opposite the throttle(s) and are red in color. A disconnect switch stops all movement of the autopilot system. An interrupt switch momentarily interrupts all movement of the autopilot system. (5) A determination that any automatic disconnects of the autopilot is adequately annunciated by an aural warning. If warning lights are utilized to supplement the aural warning, they should meet the requirements of § 23.1322. Use of a visual warning as the sole means of annunciating automatic disconnects is not considered acceptable. (6) Motion and effect of autopilot cockpit controls should conform with the requirements of §§ 23.1329(c) and 23.779. b. Malfunction Evaluations (1) Malfunction evaluation flights should be conducted with the airplane loaded at the most critical weight or the most critical c.g./weight combination. Maximum untrimmed fuel imbalance should be considered during the evaluation. If autothrottles are installed, they should be operating, and autopilot servo torque should be set to the upper tolerance limit. The simulated malfunctions should be induced at various airspeeds and altitudes throughout the airplane's airspeed and altitude envelopes. These envelopes should include the maximum operating altitude for turbocharged or high altitude airplanes, or be within 10 percent of the service ceiling for normally aspirated airplanes, and when the airplane is stabilized in the normal operational attitudes. Vertical gyro mechanical failures should not be considered. The simulated failures and subsequent corrective actions are not acceptable if they result in any of the following: (i) Loads that exceed the substantiated structural design limit loads. (ii) Acceleration that is outside the 0 to 2g envelope. The positive "g" limitation may be increased up to the positive design limit maneuvering load factor if it has been previously determined analytically that neither the simulated failure nor subsequent corrective action would result in loads beyond the design limit loads of the airplane. ### Page 190 (iii) Speeds in excess of VNE or for airplanes with an established VMO/MMO, a speed midway between VMO/MMO and the lesser of VD/MD, or the speed demonstrated under § 23.253. (iv) Deviations from the flight path including bank angle in excess of 60° or pitch attitude in excess of + 30° deviation from the attitude at which the malfunction was introduced. (v) A hazardous dynamic condition. (2) Normal Flight Malfunctions The airplane's performance should be evaluated when the effect caused by the most critical single failure condition that can be expected to occur to the system and can be detected by the pilot is induced into the autopilot system. Hidden or latent failures, in combination with detectable failures, should be considered when determining the most critical failure condition. Normal flight includes climb, cruise, and descent flight regimes with the airplane properly trimmed in all axes. Airplane configurations (combinations of gear and flaps), speeds, and attitudes should be evaluated for unsafe conditions. The more critical of the following simulated malfunctions are the following: (i) A simulated malfunction about any axis equivalent to the cumulative effect of any failure or combination of hidden failures, including manual-electric or automatic trim, if installed. (ii) The combined signals about all affected axes, if multiple axis failures can result from the malfunction of any single component. Since Amendment 3-2 to part 3 of the CAR, effective August 12, 1957, the requirements are that an autopilot system should be designed so that a single malfunction will not produce a hardover signal in more than one control axis (reference §§ 3.667(e) and 23.1329(e)). Note: A 3-second delay following pilot recognition of an autopilot system malfunction, as indicated in Item 2b(1), should be applied for normal flight malfunction evaluations. (3) Maneuvering and Approach Malfunction Maneuvering flight tests should include turns with the malfunction induced at the maximum bank angle for normal operation, up to and including the autopilot authority limits. Airplane configurations (combinations of gear and flaps), airspeeds, and altitudes should be evaluated to determine if unsafe conditions exist. Simulated malfunctions described for normal flight malfunctions as indicated in Items 3b(2)(i) ### Page 191 179 and (ii) (titled, “Normal Flight Malfunctions”) are applicable for introduction during maneuvering flight malfunction evaluation. The resultant accelerations, loads, and speeds should be within limits described for normal flight malfunctions. Malfunctions introduced during coupled approaches should not place the airplane in a hazardous attitude or an attitude that would prevent the pilot from conducting a missed approach or safe landing. Altitude losses resulting from the simulated malfunctions are to be measured accurately and presented in the Limitations Section of the Airplane Flight Manual (AFM) or approved manual material. In maneuvering and approach flight regimes, the pilot should recognize an autopilot system malfunction within 1-second. This recognition should occur as the result of a deviation of the airplane from the intended flight path, abnormal control movements, or by means of a reliable warning system that is applied. Note: Accurate measurement of altitude loss, due to an autopilot malfunction during an instrument landing approach, is essential. This altitude loss during a critical phase of flight provides the basis for establishing the minimum approach altitude during autopilot coupled approaches. The loss should be determined by measuring from the altitude at which the malfunction is induced to the lowest altitude observed during the recovery maneuver, unless instrumentation is available to measure the vertical deviation from the intended glide path to the lowest point in the recovery maneuver. In this section, Appendix 1 contains a method of measurement for approach altitude loss. Altitude losses due to malfunctions in other flight regimes, though less critical, may be determined by measuring the deviation from the flight path in a manner similar to that used for the glide slope. (4) Alternate Means of Compliance for Autopilots Incorporating Electronic Monitors/Limiting Devices Listed below are alternate means of compliance. These alternate means cite considerations for evaluating both monitors and limiting devices when functioning of such devices is necessary to prevent the airplane from exceeding the malfunction limits identified in paragraph 3b(1) of this AC. (i) Alternate Means No. 1 (A) Monitor/Limiter Inhibited With the monitor/limiter inhibited, autopilot malfunction flight testing may not cause any of the following: (1) Roll to exceed 80°. (2) Pitch to exceed +45°, -35°. ### Page 192 (3) Accelerations outside the 0g to 2.5g envelope. (4) Airspeed exceeding VNE or for an airplane having an established VMO/MMO, a speed not greater than a speed midway between VMO/MMO and the lesser of VD/MD or the speed demonstrated under § 23.253. (B) Reliability and Prerequisite Criteria (1) A fault analysis should show that the failure effect of a monitor failure, combined with an autopilot malfunction, is less than major; and (2) Pre-engagement check of the monitor is mandatory. No credit is allowed for a pilot-activated pre-engagement check unless there is a lockout device or system. (ii) Alternate Means No. 2 (A) Monitor/Limiter Inhibited With the monitor/limiter inhibited, autopilot malfunction flight testing may not cause any of the following: (1) Roll to exceed 80°. (2) Pitch to exceed +45°, -35°. (3) Accelerations outside the -0.2g to 2.5g envelope. (4) Airspeed exceeding VNE or for an airplane having an established VMO/MMO, a speed not greater than a speed midway between VMO/MMO and the lesser of VD/MD or the speed demonstrated under § 23.253. (B) Reliability and Prerequisite Criteria (1) An acceptable fault analysis showing that the failure effect of a combined monitor failure and an autopilot malfunction is less than hazardous. In addition, the failure effect of failure of a lockout device to inhibit autopilot engagement, as identified in Item (3) below, is less than major; (2) Pre-engagement check of the monitor is mandatory with either a manual or automatic activation means; and (3) Autopilot engagement is inhibited until pre-engagement check is successfully completed. ### Page 193 (iii) Alternate Means No. 3 (A) Flight tests with monitors inhibited are not required. (B) Reliability and prerequisite Criteria (1) An acceptable fault analysis showing that the failure effect of a combined monitor failure and autopilot be less than catastrophic. In addition, failure of a lockout device/system to inhibit autopilot engagement, as identified in Item (3) below, is less than hazardous; (2) Pre-engagement check of the monitor is mandatory with either a manual or automatic activation means; (3) Autopilot engagement inhibited until the pre-engagement check is successfully completed; and (4) Autopilot authority not greater than necessary to satisfactorily control the airplane. c. Recovery of Flight Control Evaluate the ability to recover flight control from the engaged autopilot system either by manual use of a quick disconnect or by physically overpowering the system. d. Performance Flights Performance evaluation tests should be conducted with the airplane loaded to its most adverse c.g. and weight condition. Autopilot performance with the servo torque values at the lowest production torque tolerance limit should be used to demonstrate safe controllability and stability. Flight tests are necessary to ensure the autopilot system performs its intended function, including all modes of operation presented for approval (reference § 23.1301). e. Single-Engine Approach For multiengine airplanes, an engine failure during a normal instrument landing system (ILS) approach should not cause a lateral deviation of the airplane from the flight path at a rate greater than 3° per second or produce hazardous attitudes. This rate should be measured and averaged over a 5-second period. If approval is sought for ILS approaches initiated with one engine inoperative, the autopilot should be capable of conducting the approach. ### Page 194 f. Airplane Flight Manual (AFM) Information The following information should be placed in the AFM (or, if the airplane does not have an AFM, it should be placed in the Pilot's Operating Handbook (POH) or presented to the pilot in the form of placards): (1) In the Operating Limitations Section, the airspeed limitations, maximum altitude for operation if different from the maximum certificated altitude of the airplane, category of ILS approaches for which approval is granted, minimum approach height, and any other applicable limitations. (2) In the Operating Procedures Section, the normal operating information, including navigation and glide slope intercept recommendations. For those autopilot systems which incorporate either monitors or limiter devices, the pre-engagement procedures and the means of indicating that the pre-engagement has been successfully completed. (3) In the Emergency Operation Procedures Section (i) A statement of the altitude loss in the cruise, climb, and descent configurations; and maneuvering flight conditions, due to possible malfunctioning of the autopilot system. (ii) A statement of the altitude loss due to malfunctions while in the approach configuration. If engine inoperative approach is approved, the altitude loss should be included. (iii) Any other procedure related to emergency procedures associated with either the autopilot or associated systems. (See Figure 1.) ### Page 195 (OCR) Altitude Loss 1. Malfunction inducement point. 2. Malfunction recognition by pilot. 3. Initiation of manual recovery action by pilot. 4. Altitude loss with no instrumentation. 5. Altitude loss with instrumentation. 3° ILS Nominal Glide Slope Airplane Wheels One Second Delay y 1:29 Slope 15.0 feet Runway Threshold ILS Glide Fath Origin FIGURE 1. ACCEPTABLE METHOD FOR DETERMINING **ALTITUDE LOSS IN APPROACH 183** ### Page 196 184 FIGURE 1¾CONTINUATION Malfunction Evaluations. The airplane should be established on the ILS glide slope and localizer in the configuration(s) with the approach speed(s) specified by the applicant for approach. Simulated automatic flight control system malfunctions should be induced at critical points along the ILS taking into consideration all design variations and their limits in automatic flight control system sensitivity and authority. The malfunctions should be induced in each axis. While the pilots may know the purpose of the flight, they should not be informed when a malfunction is to be or has been applied except through a deviation of the airplane from the intended flight path, abnormal control movements, or by means of a reliable failure warning system. After a failure, recovery should be initiated 1 second after the pilot recognizes the failure. a. A 3° glide slope should be used for these tests in order to determine the malfunction effects to be expected in service. b. For use during a coupled ILS approach, the automatic control system should not fail in such a way that it causes the airplane wheels to descend below a limit line lying below the glide slope, sloping upward at 29:l from a point 15 feet above the runway threshold. With the airplane established on the glide slope in approach configuration, at approach speed, the most critical malfunction is induced at a test altitude referenced to the runway threshold. Measure the altitude loss between the test altitude and the lowest point of the manual recovery, unless instrumentation is available to measure the vertical deviation from the intended glide path to the lowest point in the recovery maneuver. The altitude loss and the known distance to the threshold from the lowest recovery altitude are compared to the limit line. The lowest test altitude from which malfunction and manual recovery can be completed, without the airplane wheels descending below the limit line, is considered the minimum height for use of the automatic flight control system. c. Recovery from all malfunctions should be demonstrated either by overpowering or by manual use of an emergency quick disconnect device after the appropriate delay. The pilot should be able to return the airplane to its normal flight altitude under full manual control without exceeding the defined limits. ### Page 197 23.1331 Instruments using a power source Original Issue and Subsequent Paragraph (a) in § 23.1331 applies only to gyroscopic instruments, but paragraph (b) in this section applies to any instrument that depends on external power or external energy for proper operation. The requirement for two independent power sources in paragraph (b)(1) in § 23.1331 applies to either vacuum or electrically driven gyroscopic instruments. Electrical Systems When complying with paragraph (b) in this section, a single battery required for starting is acceptable if the electrical system is capable of continuous normal operation without external excitation or stability, and there is no probable failure of the battery that will adversely affect the electrical system once it is operating. However, the airplane battery cannot be accepted toward showing compliance to the power source requirements of §§ 23.1331 and 23.1351 unless the state of charge of the battery is displayed to the pilot. A single electrical bus is unacceptable for a multiengine airplane. The multiengine requirement is for two independent power sources. Therefore, an installation with a single primary power source for all flight instruments and a manually operated backup is not acceptable. a. This system could conceivably fail in such a way that all the flight instruments could be simultaneously damaged or disabled (i.e., loss of voltage regulation). This would not be remedied by switching to the backup power source. b. Also, an electrical system with a primary power source that employed a backup source with common circuitry or components is not truly independent. Standby Vacuum Systems The intended function of a standby vacuum system is to provide a second vacuum source for the gyroscopic instruments after a failure of the primary vacuum system. The standby system should either supply sufficient vacuum to maintain the accuracy and reliability of the gyroscopic instruments throughout the phases of flight, or there should be limitations on operation in the Airplane Flight Manual (AFM). When operating on the standby system, the pilot should predicate operations on other certified systems (partial panel) and use the gyroscopic instruments as an aid, provided the pilot determines these instruments give ### Page 198 186 acceptable information. Also, the pilot should not manipulate the throttle, other than for normal flight, in an attempt to control vacuum pressure within the limitations. If a second vacuum system is not required, the standby installation would be for nonrequired equipment per this rule. In addition, § 21.21(b)(2) requires there be no feature that results in an unsafe condition. To comply with these requirements, it should be shown that neither operation nor failure of the standby vacuum system interfere with the normal operation of the primary system or result in any unsafe condition. The pilot should also be kept apprised of when the standby system is in operation either by manual source selection or by red visual annunciation (§ 23.1322) if an automatic switching system is installed. To ensure that no unsafe condition will result, the standby system should be flight evaluated in each unique airplane installation. In addition, operating information, emergency procedures, and limitations should be available in an AFM Supplement, a Supplemental AFM, or placards, as appropriate. This information should meet the requirements of §§ 23.1583 and 23.1585 of this part, and it should emphasize that the standby vacuum system is for emergency use only and should not be utilized for dispatch purposes. Amendment 23-43 and Subsequent Independent Power Sources This amendment adds the requirement for independent power sources for required instruments for single-engine as well as multiengine airplanes. This was considered appropriate due to the number of single-engine airplane accidents that were attributed to the loss of power to required flight instruments. Also, the reference to “gyroscopic” was removed to include both gyroscopic and non-gyroscopic instruments since non-gyroscopic flight instruments are in use. a. Instruments that provide required flight information and use an external power source are now required to have two independent power sources. This requirement has the same intent for single-engine airplanes as for multiengine airplanes: to functionally isolate flight instruments such that any failure of one power source or instrument will not cause the complete loss of a required flight instrument function. Thus, in the case of failure of a heading instrument, that failure may not result in the loss of the proper supply of energy to the attitude indicator powered by the same source, and loss of a single power supply may not cause loss of any required instrument function. b. Ship’s batteries used in normal operations are acceptable as backup power sources only if their state of the charge can be reliably verified to the pilot. c. This regulation is not intended to apply to circuit protection devices, which are to be considered in §§ 23.1351 and 23.1357. ### Page 199 d. These changes are meant to apply to those instruments that rely on a power source and provide required flight information. Such instruments are those that provide information for direct control of flight that are required by the kinds of operation for which the airplane has been approved. Consequently, this section applies to all flight instruments required by 14 CFR part 23, § 23.1303 and part 91, § 91.205. So, instruments in airplanes limited to VFR operations that are not required for VFR would not have to comply with the requirements of § 23.1331. Exemptions would not be necessary or appropriate. e. Each independent power source must provide sufficient power for normal operations throughout the approved flight envelope of the airplane and for any operations for which the airplane is approved. For example, an IFR approved airplane must have independent power sources for the display of attitude that are not limited to altitudes below the approved service ceiling of the airplane. f. Section 23.1331(c) does not require the installation of dual alternators or vacuum systems on single engine airplanes. Other options include a dedicated battery with a 30 minute capacity for electrical instrument loads essential to continued safe flight and landing, use of differently powered types of instruments for primary and standby, or verifying the aircraft battery used for starting by a system safety analysis per § 23.1309. The last option would: (1) Require reliability and probability data for the aircraft battery that is acceptable to the FAA (applicants have not provided such data to date); and (2) Require that the state of charge be available to the pilot during all stages of flight and include a caution alert per § 23.1322 when the state of charge is less than 30 minutes. ### Page 200 23.1335 Flight director systems No policy available as of June 30, 2001. ### Page 201 23.1337 Powerplant instruments installation Original Issue and Subsequent No specific criteria have been established for the minimum orifice size for fuel and oil lines. We believe that .020 inches for fuel lines and .060 inches for oil lines can be accepted (per Air Force Systems Command Manual 80-1, Part C, Chapter 5, paragraphs 3.1.1.3.7 and 3.1.2.3.3). See AC 23-16, Powerplant Guide for Certification of Part 23 Airplanes, § 23.959, for unusable fuel test procedures for guidance on § 23.1337(b)(1). Changes to total fuel quantity by incorporation of a fuel tank filler connection (§ 23.973) outboard of the existing connection will require changing the fuel quantity indicator to indicate the new quantity of fuel. The new indicator should meet the accuracy as specified in TSO- C55, Fuel and Oil Quality Instruments (For Reciprocating Engine Aircraft), or MIL-G- 9798. The rule requires that fuel quantity be calibrated as zero when only unusable fuel is left in the tank. Some fuel tank system designs can result in a lowest reading obtainable in level flight being greater than the unusable fuel supply. In this case, an equivalent level of safety is acceptable by placing a red radial at the “lowest reading obtainable in level flight,” and mounting a placard stating the amount of usable fuel remaining at the red radial. Fuel quantity indicators are also governed by § 23.1301 as are all 14 CFR part 23, Subpart F appliances. This regulation requires the installed indicators function as designed and not create a hazard in their operation. This precludes indicators that read higher than the actual fuel level since this would constitute a hazard. 14 CFR part 23 does not require an applicant to install a TSO’d fuel quantity indicator, but when installed in a reciprocating engine airplane and produced under TSO-C55 authorization, the allowable error of the indicator is no more than three percent (3 percent) of full scale. It is believed that ground and cruise attitude(s) are the minimum attitudes required for operation within the 3 percent tolerance. Evaluation of the gauge function throughout the normal and expected operation of the airplane is needed to assess gauge indications outside the ground and cruise attitude conditions. The purpose of this evaluation is to establish that the ground and cruise attitude(s) gauge tolerance is functional and adequate for the airplane. Amendment 23-18 and Subsequent See AC 23-8A, Flight Test Guide for Certification of Part 23 Airplanes, for guidance on fuel quantity indicators and auxiliary tanks. ### Page 202 **ELECTRICAL SYSTEMS AND EQUIPMENT** 23.1351 General Original Issue This rule does not allow a failure or malfunction of any electrical power source to impair the ability of any other source to supply essential circuits. Amendment 23-7 and Subsequent This amendment allows one exception to the original rule. This exception would allow loss of an alternator that is dependent on a battery for initial excitation or stabilization when that battery has failed. This exception was adopted under the premise that the advantages of having a battery connected for initial excitation or stabilization for alternators needing it outweighed the consequences of that battery failing. This exception is only applicable to alternator installations that need a battery. Loss of an alternator due to battery failure was considered of no greater consequence than the intrinsic failure of the alternator itself. In the case of single-engine airplanes, loss of the battery and alternator would result in the loss of the electrical system, which would be no worse than other single failures (i.e., shorts to ground, conductor failure, etc.) that would also result in loss of the electrical system. Wire meeting MIL-W-5086 has been removed from the listing of approved wire in AC 43.13-1B due to its flammability characteristics, corrosive vapors, and toxic gases of PVC insulation. See AC 43.13-1B, Acceptable Methods, Techniques, and PracticesAircraft Inspection and Repair, Section 3, for allowable wire in airplane manufacture and alteration. Wire that is not listed in AC 43.13-1B should be shown by tests and analyses to meet the airworthiness requirements of § 23.863, 23.1351, 23.1359, and 23.1365 as applicable. JAA ACJ 23.1351(a)(2) is acceptable for FAA certification. JAA ACJ 23.1351(b)(5)(iv) is acceptable for FAA certification to 14 CFR part 23, § 23.1351(b)(4)(iv). ### Page 203 23.1353 Storage battery design and installation Original Issue and Subsequent See AC 43.13-1B, Acceptable Methods, Techniques, and PracticesAircraft Inspection and Repair, Section 8, for battery installation guidance. Replacement batteries would require Parts Manufacturer Approval (PMA), unless exempted under the provisions of part 21, § 21.303(b), whether the replacements are lead-acid or nickel-cadmium. The airworthiness standards of 14 CFR part 23, §§ 23.1301, 23.1309, 23.1351, and 23.1353 should be considered for replacement battery installations. Amendment 23-49 and Subsequent Thirty Minutes of Electrical Power Requirement by § 23.1353(h) This guidance regards the 30 minutes of electrical power requirement under Title 14 Code of Federal Regulations (14 CFR) part 23 incorporated by Amendment 23-49 into § 23.1353(h). This guidance only addresses the requirement of § 23.1353(h) and not the electrical power requirements that an airplane can operate safely in VFR conditions under § 23.1351(g) or the electrical power sources requirements under § 135.163. The requirements of § 23.1353(h) are as follows: “In the event of a complete loss of the primary electrical power generating system, the battery must be capable of providing at least 30 minutes of electrical power to those loads that are essential to continued safe flight and landing. The 30 minute time period includes the time needed for the pilots to recognize the loss of generated power and take appropriate load shedding action.” Per previous guidance on this issue, the airplane’s primary electrical power includes the airplane’s electrical generation system and the airplane’s starter battery when only one battery is installed. The battery for the 30-minute criteria, therefore, should be an independent power source from the airplane’s starter battery. If adequate monitoring and procedures are incorporated so the pilot knows that the airplane’s starter battery meets the 30 minute criteria after an engine start and during all other operations, an equivalent level of safety finding may be an acceptable method for using the airplane’s starter battery. Please refer to the guidance in AC 23.1309-1C, Equipment, Systems, and Installations in Part 23 Airplanes, for determining the loads that are essential to continued safe flight and landing. Continued safe flight and landing is defined as follows: ### Page 204 192 “This phrase means that the airplane is capable of continued controlled flight and landing, possibly using emergency procedures, without requiring exceptional pilot skill or strength. Upon landing, some airplane damage may occur as a result of a failure condition.” The 30-minute power bus should include all systems that could cause a catastrophic failure condition under the § 23.1309, Failure Hazard Assessment. In some cases, it may not be practical to include all systems on the 30-minute power bus that could cause a catastrophic failure condition. For example, systems with large heating loads for ice protection may not be included on the 30 minute electrical power bus; however, the possible hazards that could cause catastrophic failure conditions should be minimized. To minimize the hazard is to reduce, lessen, or diminish to the least practical amount with current technology and materials. The least practical amount is that point at which the effort to further reduce a hazard significantly exceeds any benefit in terms of safety derived from that reduction. Additional efforts would not result in any significant improvements to safety and would inappropriately add to the cost of the product. Assuming operations under IFR conditions for part 91 or 135 operations, the following systems should be included on the 30-minute power bus: • If needed to comply with § 23.1325, one airspeed indicator and altimeter with a heated pitot tube and heated static pressure source; • The magnetic compass, and any display necessary for continued safe flight and landing, sufficiently illuminated for night operation; • One navigation system installation appropriate to the ground facilities to be used; • One communication installation system; • One gyroscopic pitch and bank indicator; • One clock; • Any display for the powerplant parameter necessary for continued safe flight and landing; and • Any electrical loads unique for the airplane characteristics and needed for continued safe flight and landing for the intended operations. Tests and analyses should be considered for determining the rated operating capacity of the battery, the normal service life, and the continued airworthiness requirement of § 23.1529. For these tests and analyses, the following should be established: ### Page 205 (1) For the operating capacity: the discharge rate, temperature, end-point voltage, etc.; and (2) For the airworthiness requirement: the inspection schedule, useful battery life, end-of-life, etc. JAA ACJ 23.1353(h) is acceptable for FAA certification. ### Page 206 23.1357 Circuit protective devices Original Issue and Subsequent This rule allows only one essential circuit on one circuit protective device. The intent of the rule is met for installations that integrate position and anticollision lights on one wire when no single failure can cause the loss of any of the following: (a) More than all anticollision lights. (b) More than all position lights. (c) More than one position light and the anticollision light adjacent to it. The requirement in § 23.1357(e) applies to fuses for all circuits, not just essential circuits. Although spare fuses for non-essential systems and equipment do not have to be resettable in flight. The phrase “essential to safe operation,” as used in part 135, Appendix A, paragraph 64, and the phrase “essential to flight safety” in § 23.1357(b) have the same meaning as “essential to safety in flight” in § 23.1357(d) and “essential to flight safety” in § 23.1357(b). All of these phrases are descriptive of equipment installed in order to comply with the airworthiness or operational requirements. The FAA recognizes that some required circuit protection devices are associated with circuits that can have no significant impact on safety in flight. Therefore, the responsible Aircraft Certification Office (ACO), in conjunction with the applicant, should identify which circuits and circuit protection devices are essential to safety in flight. The identified circuits should comply with § 23.1357(d) regarding the pilot’s ability to reset them in flight. Fuel quantity indicators are required by regulation but are not essential to safety in flight. The loss of an indicator will increase the pilot’s workload, but it will not, in itself, cause either a loss or a forced landing of the airplane. It is acceptable, therefore, to have a protected circuit that includes more than one required indicator, which includes a fuel quantity indicator. Good design practice would include placing multiple indicators of the same parameter on separate protected circuits, and we encourage the use of warnings for low fuel, high oil temperature, etc., to mitigate the effects of loss of indication. The intent of § 23.1357(b) is that, in the case of an essential load, its individual circuit should be the only load on an individual circuit protection device. This intent was explicitly stated in Amendment 23-20. ### Page 207 195 For part 23 applications, the definitions of a switch and a circuit breaker are as follows: a switch is a device for opening and closing or for changing the connection of a circuit; a circuit breaker is a device designed to open and close a circuit by non-automatic means and to open the circuit automatically at a predetermined overload of current, without injury to itself when properly applied within its rating. Consequently, circuit breakers used for operational functions are not acceptable in that they are not performing their intended function, which is protection against overloads. Circuit breakers, even those suitable for frequent operation, should not be used as a switch to perform procedural functions. A combination switch/circuit breaker is a device which can perform both as a switch for opening and closing a circuit as well as a circuit breaker, automatically opening the circuit at a predetermined overload current. ### Page 208 23.1359 Electrical system fire protection See guidance in Section 23.853 of this AC. ### Page 209 23.1361 Master switch arrangement No policy available as of June 30, 2001. ### Page 210 23.1365 Electric cables and equipment Original Issue and Subsequent Section 23.1365 requires that each cable that would overheat in a circuit overload or malfunction be at least flame resistant and not emit dangerous quantities of toxic fumes. The compliance methods for the flame resistance requirement are in AC 23-2, Flammability Tests. To aid in meeting the toxic fume requirement, the FAA has removed MIL-W-5086 wire from the listing of approved wires in AC 43.13-1B, Acceptable Methods, Techniques, and PracticesAircraft Inspection and Repair. See guidance in Section 23.853 of this AC for flammability. Amendment 23-14 and Subsequent The flame resistance and toxic fume requirements are applicable to equipment associated with the cable as well as the cable itself. ### Page 211 23.1367 Switches Original Issue and Subsequent Switches are required to be labeled as to operation and the circuit controlled. A switch that operates by a push once for ON and once for OFF should be labeled “PUSH OFF/ON.” Switches are also required to be accessible to the flight crew. The intent of this rule is that those switches that are installed in the cockpit should be accessible to a flight crew member if manual operation is necessary for safety of flight. ### Page 212 **200 LIGHTS** 23.1381 Instrument lights No policy available as of June 30, 2001. ### Page 213 23.1383 Taxi and landing lights No policy available as of June 30, 2001. ### Page 214 23.1385 Position light system installation Original Issue and Subsequent Guidance on light measurements can be found in AC 20-74, Aircraft Position and Anticollision Light Measurements. Additional guidance on position lights can be found in AC 20-30B, Aircraft Position Light and Anticollision Light Installation. Guidance on flame resistance can be found in AC 23-2, Flammability Tests. The intent of the rule in § 23.1357 is met for installations that integrate position and anticollision lights on one wire when no single failure can cause the loss of any of the following: (a) More than all anticollision lights. (b) More than all position lights. (c) More than one position light and the anticollision light adjacent to it. Position lights are not required for airplanes limited to Day VFR operation (placarded for VFR Day). If approved for Night VFR or IFR, then position lights are required per §§ 23.1385 through 23.1395. They should be listed on the kinds of operation equipment list (§ 23.1559(b)) and included in the Limitations Section of the Airplane Flight Manual (§ 23.1583(h)). ### Page 215 23.1387 Position light system dihedral angles No policy available as of June 30, 2001. ### Page 216 23.1389 Position light distribution and intensities No policy available as of June 30, 2001. ### Page 217 23.1391 Minimum intensities in the horizontal plane of position lights No policy available as of June 30, 2001. ### Page 218 23.1393 Minimum intensities in any vertical plane of position lights No policy available as of June 30, 2001. ### Page 219 23.1395 Maximum intensities in overlapping beams of position lights No policy available as of June 30, 2001. ### Page 220 23.1397 Color specifications Original Issue and Subsequent See AC 20-74, Aircraft Position and Anticollision Light Measurements, for guidance on color measurements. ### Page 221 23.1399 Riding light No policy available as of June 30, 2001. ### Page 222 23.1401 Anticollision light system Original Issue and Subsequent If certification for night operation is requested, an anticollision light system, per this section, is required. See AC 20-74, Aircraft Position and Anticollision Light Measurements, for guidance on anticollision light measurements. The flash rate of supplemental lights does not have to be applied to the anticollision light flash rate, but these lights should be checked to verify there is no unsafe condition associated with their use. Amendment 23-11 and Subsequent There is no restriction on mixing aviation red and aviation white anticollision lights on the same airplane. Likewise, there is no restriction on the ratio of red to white provided that the light displayed in any one direction is either aviation red or aviation white. Some white supplementary lights have been presented for certification as anticollision lights. The visible limit of such lights may converge at some point forward and aft of the airplane such that from this point to the airplane neither light is visible. The maximum allowable distance to such convergence is 1,200 feet. The regulations (§ 23.1397) require that aviation white’s “X” coordinate be no less than 0.300 and no greater than 0.540 (International Civil Aviation Organization (ICAO) Annex 8 requirement). Xenon flash tubes can exceed the “X” limit for some energy levels (20 to 40 joule range). For the function of an anticollision light, an occasional excursion beyond the 0.300 limit would not adversely affect safety or the performance of the intended function. We have been advised by the National Bureau of Standards (NBS) that the measurement accuracy of the “X” value of chromaticity coordinates includes an error tolerance of plus or minus 0.008. It was not envisioned that filtering would be required on Xenon flash tubes to meet the aviation white limits since the color can be effectively limited by capacitor circuitry to control the energy level of individual flashes. The maximum joules per flash should be such that the 0.300 will not be exceeded more than 68 percent of the time and 0.292 will not be exceeded 99.7 percent of the time (3 sigma), which includes the measurement error tolerance suggested by NBS. ### Page 223 **SAFETY EQUIPMENT** 23.1411 General No policy available as of June 30, 2001. ### Page 224 23.1413 Safety Belts and Harnesses [Removed] See guidance for Section 23.785 in this AC. ### Page 225 23.1415 Ditching equipment No policy available as of June 30, 2001. ### Page 226 23.1416 Pneumatic de-icer boot system Policy is available in AC 23.1419-2A. ### Page 227 23.1419 Ice protection The § 23.1419 guidance in this advisory circular is focused on the ice protection systems at the component and system level. For information pertaining to airplane performance and handling in icing conditions please refer to AC 23.1419-2A, Certification of Part 23 Airplanes for Flight In Icing Conditions, and AC 20-73, Aircraft Ice Protection. Original Issue and Subsequent Icing Policy for Small Airplanes The Civil Air Regulations (CAR) 3 airplanes and 14 CFR part 23 airplanes with a certification basis preceding Amendment 23-14, are permitted to fly in known icing conditions if their Type Certificates do not include a prohibition against this operation. Some of these airplanes may be placarded against flight into known icing because they lack deice/anti-ice equipment specified in the type design data. Installation of equipment required for icing approval per type design data is justification for removing the placard. However, the part 91 and part 135 operating rules in icing have limitations for these aircraft. If the above airplanes’ type design data prohibits flight into known icing, then these airplanes can be approved for flight into known icing only if compliance is shown to part 23, § 23.1419, Amendment 23-14, and subsequent. Amendment 23-14 and Subsequent To certificate a single-engine airplane for flight in icing conditions, part 21, § 21.101 would require the same criteria to be applied as in VFR, IFR, Day and Night flight, which is to keep the airplane in the air and flying even if performance is compromised somewhat. This may require redundancy in ice protection system components to minimize hazards to the airplane in the event of a probable malfunction or failure. There is no requirement or allowance for making adjustments in the icing certification program for the frequency of encountering icing conditions. A probability of one is to be used for encountering discrete environmental conditions such as instrument meteorological conditions. Icing conditions are environmental conditions, and an encounter frequency of less than one for compliance with §§ 23.1093 and 23.1419 is not appropriate. Amendment 23-43 and Subsequent See AC's 23.1419-2, 23-143-1, 23-8 and 20-73. ### Page 228 ### Page 229 **MISCELLANEOUS EQUIPMENT** 23.1431 Electronic equipment Original Issue and Subsequent Mercury cell battery packs for use in Emergency Locator Transmitters (ELT) should be manufactured by controlled processes. Service experience has shown that ELT mercury cell battery packs fabricated by individuals without a controlled process can result in the following: a. Degradation of the cell seal causing leaks and a shorter shelf life. b. Creation of internal shorts. c. Internal corrosion. d. Creation of highly explosive mercury fulminate. The possibility of adverse interaction between communication and navigation equipment should be evaluated. Momentary indicator deflection or flicker is acceptable. However, loss of a required function due to interaction of assignable frequencies in the National Airspace System is not acceptable. Interim guidance for TCAS I installations is given in AC 20-TCAS, Airworthiness Approval and Operational Use of Traffic Alert and Collision Avoidance System (TCAS I) (in draft). Guidance for TCAS II installations is given in AC 20-131A, Airworthiness and Operational Approval of Traffic Alert and Collision Avoidance Systems (TCAS II) and Mode S Transponders. Automatic NAVAID selection tuning (Auto-tune) of Very High Frequency Omnirange Station/Distance Measuring Equipment (VOR/DME) for flight management or multisensor navigation systems is designed to enhance the navigation accuracy for enroute flight. Under certain conditions, and depending on the particular implementation, the auto-tune function can cause a hazard if auto-tune remains operative during VOR and ILS operations. In this case, automatic selection of a NAVAID different than that wanted by the flight crew is a possibility. Visual cues indicating the auto-tune is still active may be quite subtle and may go unnoticed during a high workload period. If the auto-tune NAVAID is reasonably in line with the projected track, the anomaly can go undetected—causing the airplane to fly an erroneous track based on the auto-tune NAVAID. This may occur either when steering manually or when the flight guidance system has been engaged. System installations that ### Page 230 218 employ auto-tune should be mechanized in a manner that addresses these safety issues. An acceptable method of auto-tune implementation is to automatically inhibit the auto-tune feature when a navigation function other than the one utilizing auto-tune has been selected for display on the Horizontal Situation Indicator/Electronic Horizontal Situation Indicator (HSI/EHSI). Moving Map Displays (MMD) used for primary command guidance during IFR flight should be evaluated with the particular navigation receiver (GPS, LORAN-C, etc.) to be used. It should also be restricted to use with that particular type of receiver on that particular airplane. If a separate command or deviation indicator is used to certify the system for IFR use, the MMD should be placarded “For Reference Only” and used only if it can be shown that failure of the MMD would not fail the navigation system. An MMD to be used for and placarded for “VFR Only” guidance would need verification that it performs its intended function when used with a particular navigation receiver. It could then be used on any other airplane with the same type navigation receiver. JAA ACJ 23.1431(e) is acceptable for FAA certification. ### Page 231 23.1435 Hydraulic systems No policy available as of June 30, 2001. ### Page 232 23.1437 Accessories for multiengine airplanes No policy available as of June 30, 2001. ### Page 233 23.1438 Pressurization and pneumatic systems Amendment 23-20 and Subsequent Burst and proof tests are required at multiples of the Maximum Normal Operating Pressure. Temperature effects are not required as part of these tests if pressurization and pneumatic system elements are constructed of materials that can withstand the operational pressures and temperatures. If sufficiently overstrength material elements are part of a design, then reduced material strength due to temperature variations is not a concern and testing may be performed at ambient temperature. The “sufficiently overstrength” determination must consider temperature cycles/extremes for materials that may exhibit structural property phenomenon and structural strength changes due to temperature changes. Certain composite materials may be characteristic of adverse performance in certain temperature environments. Furthermore, the “sufficiently overstrength” determination must consider both the duct system and its attach/clamping/stabilization devices. This advice is consistent with §§ 23.1301, Function and Installation; 23.307, Proof of Structure; and 23.603(a), Materials and Workmanship. ### Page 234 23.1441 Oxygen equipment and supply Amendment 23-9 and Subsequent Plastic lines (nylon, polyvinyl chloride (PVC) and Teflon) are not acceptable for use in continuously pressurized, non-portable oxygen systems. Plastic lines can be used in non-portable oxygen systems that are pressurized only when cabin decompression occurs with the following precautions: a. Swaged metal type end fittings should be used to prevent leakage from cold flow. b. Lines should be protected from abrasion by use of a reinforcing sleeve of fabric braid. c. Lines should be routed away from areas where they might be subjected to elevated temperatures, electrical arcing (relays and switches), and flammable fluids. d. Refer to AC 43.13-2A, Acceptable Methods, Techniques, and Practices Aircraft Alterations, Chapter 6; and NASA Safety Standard for Oxygen and Oxygen Systems, NSS 1740.15, dated January 1996, for additional guidance material. Design and install oxygen tubes/hoses in such a manner that the hoses are stable during all phases of flight. It is not adequate to require the applicant to design tube/hose routings and chaffing protection without ensuring (by design) that the installed tubes/hoses will not vibrate/flap in a manner that would defeat the chaffing prevention wrap, or to impact other system elements such as sensors, wires, mechanisms. Part 23 is unique in that it allows oxygen system requirements to be met with portable systems. For those portable systems, information should be provided to the flight crew in the form of limitations stating which portable system is approved, which components constitute the system, and any operating limitations. Part 23 airplanes may be certified with or without an oxygen system. The necessity for supplemental oxygen is a function of the operational altitude not the airplane design. Therefore, the requirements for when supplemental oxygen is required can be found in General Operating and Flight Rules. If installed, the system should meet the following part 23 airworthiness requirements: (a) §§ 23.1441 through 23.1449 (and § 23.1450 if chemical oxygen generators are used), and (b) it may be a basic part of the airplane or a portable system. Section 23.1525 requires the airplane operational limits be established in accordance with the installed equipment or lack thereof. If an airplane is delivered without ### Page 235 223 an oxygen system, its Airplane Flight Manual should have a limitation or there should be a placard prohibiting flight above 14,000 feet mean sea level (MSL). Amendment 23-43 and Subsequent Under the previous amendment, certification of an oxygen system was at the discretion of the applicant. Under this amendment, an oxygen system “must be provided” if an airplane is certificated to operate at altitudes where the operational rules require oxygen use. This rule is applicable to any airplane with a certification basis of Amendment 23-43 or later. The rule permits portable systems to be designed by a type certificate holder subject to actions to prevent the portable system from being a hazard by not being properly secured in an emergency landing (per § 23.561). The precise requirement regarding portable systems should be determined in reference to the specifics of the airplane type (i.e., its service ceiling, number of seats, etc.). The rulemaking history does not state a safety deficiency in current methods of providing supplemental oxygen by the use of operator-provided, portable, constant flow oxygen systems either attached to the seatback or restrained by a seatbelt. If a type certificate holder does not plan to provide an oxygen system, it must specify a maximum weight and size of portable oxygen bottle and a specific means of restraining the bottle in an emergency landing per § 23.561 either by attachment to a seatback or a seatbelt. This could be commercially available or produced by the type certificate holder and available with the airplane as standard equipment or as an option. If an applicant does not provide the oxygen system or restraint information for portable bottles, the airplane will be limited to operational altitudes where oxygen is not required. ### Page 236 23.1443 Minimum mass flow of supplemental oxygen Original Issue and Subsequent When there is full compliance to this regulation, there is no need to consider the probability of a pressurization failure or to require an immediate descent in altitude in the event of a failure. Compliance should include the consideration of a rapid/explosive decompression to ambient pressure with a pilot recognition and reaction time of 17 seconds to initiate a descent. (a) The airplane may be altitude-limited to meet this requirement, or (b) The applicant may provide an equivalent level of safety finding. ### Page 237 23.1445 Oxygen distribution system Amendment 23-43 and Subsequent The guidance in this AC for § 23.1441, Amendment 23-9 and subsequent, for plastic lines is applicable to this regulation. ### Page 238 23.1447 Equipment standards for oxygen dispensing units Amendment 23-9 and Subsequent The Small Airplane Directorate was recently asked for applicable policy regarding the requirements for (1) operation above 25,000 feet, dispensing units be immediately available to each occupant wherever seated, and for (2) operation above 30,000 feet, dispensing units be automatically presented to each occupant before the cabin pressure altitude exceeds 15,000 feet. Item (1) above came into part 23 in Amendment 23-9 and item (2) above came in at Amendment 23-20. These rules were based on seats that were fixed in location and orientation. The question arises from the use of swiveling and tracked seats where it is possible for some occupants to move their seat into an orientation where an automatically presented dispensing unit will not be in view. The following guidance is applicable to an airplane with a certification basis of Amendment 23-9 or subsequent. This guidance is taken from Advisory Circular 25-17, Section 25.1447, where the issue of tracked and swiveled seats has been addressed. Automatic presentation is acceptable if the dispensing unit (mask) is presented in front of the eyes when the person’s head is resting on the seat back cushion with the seat in any position, such as upright, reclined, swiveled or tracked. The mask need not be presented in front of all persons if there is sufficient “crowd awareness,” i.e., the vast majority have proper presentation and the others can readily see that the masks have been presented. These latter persons should have a preflight briefing clearly showing them the location of their mask. The mask should be reachable with the seat belt fastened. In some seating arrangements, such as executive interiors, the various seat positions result in many different combinations of person groupings. Each combination should have an adequate number of masks reachable by every person. Consideration should be given to minimizing the likelihood of persons taking the wrong masks, thus depriving another person of their mask. If the mask must be pulled to initiate oxygen flow, the mask should be presented so that the person must pull the mask to don it. The fifth percentile female and ninety-fifth percentile male should be considered. For such as sleeper seats, bunks or lavatories, a streamer of webbing attached to the mask is acceptable to enable the person to pull the mask down to them. Amendment 23-30 and Subsequent This amendment allows the use of nasal cannulas for operation up to an altitude of 18,000 feet Mean Sea Level (MSL). These are simple devices with no known service problems, and the FAA has not developed a design standard for them. ### Page 239 227 Section 23.1447(e) requires that oxygen masks be automatically presented to each occupant before the cabin pressure exceeds 15,000 feet for airplanes certificated for operation above 30,000 feet MSL. So, just before the cabin pressure altitude exceeds 15,000 feet, the oxygen mask should fall down automatically and present itself to a 95th percentile human occupant at mouth level within the visual periphery. All the occupant should have to do is pull the mask from the hanging position, don the mask, and start breathing. ### Page 240 23.1449 Means for determining use of oxygen No policy available as of June 30, 2001. ### Page 241 23.1450 Chemical oxygen generators No policy available as of June 30, 2001. ### Page 242 23.1451 Fire protection for oxygen equipment No policy available as of June 30, 2001. ### Page 243 23.1453 Protection of oxygen equipment from rupture No policy available as of June 30, 2001. ### Page 244 23.1457 Cockpit voice recorders No FAA policy is available as of June 30, 2001. JAA ACJ 23.1459(b) is acceptable for FAA certification. ### Page 245 23.1459 Flight recorders No FAA policy is available as of June 30, 2001. JAA ACJ 23.1459(b) is acceptable for FAA certification. ### Page 246 23.1461 Equipment containing high energy rotors Amendment 23-20 and Subsequent This regulation requires that equipment containing high-energy rotors meet § 23.1461(b), (c) or (d). An acceptable means of compliance to § 23.1461 is given in AC 20-128A, Design Considerations for Minimizing Hazards Caused by Uncontained Turbine Engine and Auxiliary Power Unit Rotor Failure. --- OpenAirDocs.org is an independent research project of [Contrail Labs](https://contraillabs.com/). It is not affiliated with the FAA or the NTSB. Source documents are U.S. government works in the public domain.