AC 23-17B 

SYSTEMS AND EQUIPMENT GUIDE 

FOR CERTIFICATION OF PART 23 AIRPLANES 

AND AIRSHIPS 

April 12, 2005 

For sale by the Superintendent of Documents, U.S. Government Printing Office, Washington, DC 20401

Initiated by ACE-100 

04/12/05 AC 23-17B 

FOREWORD 

This Advisory Circular (AC) sets forth an acceptable 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 and airships. The policy in this 

AC is considered applicable for airship projects; however, the certifying office should only use 

specific applicability and requirements if they are determined to be reasonable, applicable and 

relevant to the airship project. This AC applies to Subpart D from § 23.671 and Subpart F. This 

AC both consolidates existing policy documents, and certain ACs that cover specific paragraphs 

of the regulations, into a single document and adds new guidance. 

This document is intended to provide guidance for the original issue of part 23 and the 

various amendments. This version of the AC covers policy available through 

September 30, 2003. Policy that became available after September 30, 2003, will be 

covered in future revisions to the AC. 

S/ 

David R. Showers 

Acting Manager, Small Airplane Directorate 

Aircraft Certification Service 

i (ii blank) 

04/12/05 AC 23-17B 

CONTENTS

 PAGE 

1. PURPOSE....................................................................................................................... 1 

2. CANCELLATION or SUPERSEDING ......................................................................... 1 

3. BACKGROUND ............................................................................................................ 1 

4. APPLICABILITY........................................................................................................... 1 

5. PARAGRAPHS KEYED TO PART 23 ......................................................................... 1 

6. RELATED PUBLICATIONS ........................................................................................ 1 

7. RELATED 14 CFR, PART 23 REGULATIONS........................................................... 7 

Subpart D—Design and Construction 

CONTROL SYSTEMS

Section 

23.671 General........................................................................................................... 9 

23.672 Stability augmentation and automatic and power-operated systems............. 10 

23.673 Primary flight controls................................................................................... 11 

23.675 Stops............................................................................................................... 12 

23.677 Trim systems.................................................................................................. 13 

23.679 Control system locks...................................................................................... 16 

23.681 Limit load static tests..................................................................................... 18 

23.683 Operation tests ............................................................................................... 19 

23.685 Control system details.................................................................................... 28 

23.687 Spring devices................................................................................................ 29 

23.689 Cable systems................................................................................................. 30 

23.691 Artificial stall barrier system ......................................................................... 31 

23.693 Joints.............................................................................................................. 34 

23.697 Wing flap controls.......................................................................................... 35 

23.699 Wing flap position indicator .......................................................................... 36 

23.701 Flap interconnection ...................................................................................... 37 

23.703 Takeoff warning system................................................................................. 42 

iii 
AC 23-17B 04/12/05 

LANDING GEAR 

Section PAGE

23.721 General........................................................................................................... 43 

23.723 Shock absorption tests ................................................................................... 44 

23.725 Limit drop tests.............................................................................................. 45 

23.726 Ground load dynamic tests ............................................................................ 46 

23.727 Reserve energy absorption drop test.............................................................. 47 

23.729 Landing gear extension and retraction system............................................... 48 

23.731 Wheels............................................................................................................ 53 

23.733 Tires............................................................................................................... 54 

23.735 Brakes ............................................................................................................ 67 

23.737 Skis................................................................................................................. 85

23.745 Nose/tail wheel steering................................................................................. 86 

FLOATS AND HULLS

23.751 Main float buoyancy ...................................................................................... 87 

23.753 Main float design ........................................................................................... 88 

23.755 Hulls............................................................................................................... 89 

23.757 Auxiliary floats.............................................................................................. 90 

PERSONNEL AND CARGO ACCOMMODATIONS 

23.771 Pilot compartment.......................................................................................... 91 

23.773 Pilot compartment view................................................................................. 92 

23.775 Windshields and windows ............................................................................. 94 

23.777 Cockpit controls............................................................................................. 97 

23.779 Motion and effect of cockpit controls............................................................ 100 

23.781 Cockpit control knob shape ........................................................................... 101 

23.783 Doors.............................................................................................................. 102 

23.785 Seats, berths, litters, safety belts, and shoulder harnesses............................. 105 

23.787 Baggage and cargo compartments................................................................. 118 

23.791 Passenger information signs .......................................................................... 122 

23.803 Emergency evacuation................................................................................... 123 

23.805 Flight crew emergency exits.......................................................................... 124 

23.807 Emergency exits............................................................................................. 125 

23.811 Emergency exit marking................................................................................ 135 

23.812 Emergency lighting........................................................................................ 137 

23.813 Emergency exit access................................................................................... 138 

23.815 Width of aisle................................................................................................. 139 

23.831 Ventilation ..................................................................................................... 140 

iv 
04/12/05 AC 23-17B 

PRESSURIZATION

Section PAGE

23.841 Pressurized cabins.......................................................................................... 144 

23.843 Pressurization tests......................................................................................... 149 

FIRE PROTECTION

23.851 Fire extinguishers........................................................................................... 150 

23.853 Passenger and crew compartment interiors ................................................... 151 

23.855 Cargo and baggage compartment fire protection........................................... 156 

23.859 Combustion heater fire protection ................................................................. 158 

23.863 Flammable fluid fire protection ..................................................................... 159 

23.865 Fire protection of flight controls, engine mounts, and other 

flight structure................................................................................................ 160 

ELECTRICAL BONDING AND LIGHTNING PROTECTION

23.867 Electrical bonding and protection against lightning and 

 static electricity.............................................................................................. 163 

MISCELLANEOUS 

23.871 Leveling means.............................................................................................. 165 

Subpart F—Equipment 

GENERAL 

23.1301 Function and installation................................................................................ 166 

23.1303 Flight and navigation instruments.................................................................. 172 

23.1305 Powerplant instruments.................................................................................. 180 

23.1307 Miscellaneous equipment .............................................................................. 196 

23.1309 Equipment, systems, and installations........................................................... 198 

INSTRUMENTS: INSTALLATION

23.1311 Electronic display instrument systems........................................................... 216 

23.1321 Arrangement and visibility ............................................................................ 222 

23.1322 Warning, caution, and advisory lights........................................................... 225 

23.1323 Airspeed indicating system............................................................................ 227 

23.1325 Static pressure system.................................................................................... 230 

v 
AC 23-17B 04/12/05 

Section PAGE

23.1326 Pitot heat indication systems.......................................................................... 235 

23.1327 Magnetic direction indicator.......................................................................... 237 

23.1329 Automatic pilot system .................................................................................. 240 

23.1331 Instruments using a power source.................................................................. 253 

23.1335 Flight director systems................................................................................... 257 

23.1337 Powerplant instruments installation............................................................... 258 

ELECTRICAL SYSTEMS AND EQUIPMENT

23.1351 General........................................................................................................... 261 

23.1353 Storage battery design and installation .......................................................... 265 

23.1357 Circuit protective devices .............................................................................. 270 

23.1359 Electrical system fire protection .................................................................... 272 

23.1361 Master switch arrangement............................................................................ 273 

23.1365 Electric cables and equipment ....................................................................... 275 

23.1367 Switches......................................................................................................... 277 

LIGHTS

23.1381 Instrument lights............................................................................................ 278 

23.1383 Taxi and landing lights .................................................................................. 279 

23.1385 Position light system installation ................................................................... 280 

23.1387 Position light system dihedral angles............................................................. 282 

23.1389 Position light distribution and intensities ...................................................... 284 

23.1391 Minimum intensities in the horizontal plane of position lights..................... 285 

23.1393 Minimum intensities in any vertical plane of position lights ........................ 286 

23.1395 Maximum intensities in overlapping beams of position lights...................... 287 

23.1397 Color specifications ....................................................................................... 288 

23.1399 Riding light .................................................................................................... 289 

23.1401 Anticollision light system.............................................................................. 290 

SAFETY EQUIPMENT 

23.1411 General........................................................................................................... 294 

23.1413 Safety Belts and Harnesses [Removed]......................................................... 295 

23.1415 Ditching equipment........................................................................................ 296 

23.1416 Pneumatic de-icer boot system ...................................................................... 297 

23.1419 Ice protection ................................................................................................. 298 

vi 
04/12/05 AC 23-17B 

MISCELLANEOUS EQUIPMENT

Section PAGE

23.1431 Electronic equipment ..................................................................................... 300 

23.1435 Hydraulic systems.......................................................................................... 304 

23.1437 Accessories for multiengine airplanes........................................................... 306 

23.1438 Pressurization and pneumatic systems........................................................... 307 

23.1441 Oxygen equipment and supply....................................................................... 309 

23.1443 Minimum mass flow of supplemental oxygen............................................... 313 

23.1445 Oxygen distribution system ........................................................................... 316 

23.1447 Equipment standards for oxygen dispensing units ........................................ 317 

23.1449 Means for determining use of oxygen ........................................................... 321 

23.1450 Chemical oxygen generators.......................................................................... 322 

23.1451 Fire protection for oxygen equipment ........................................................... 323 

23.1453 Protection of oxygen equipment from rupture............................................... 324 

23.1457 Cockpit voice recorders................................................................................. 325 

23.1459 Flight recorders.............................................................................................. 326 

23.1461 Equipment containing high energy rotors...................................................... 327 

vii (and viii) 

04/12/05 AC 23-17B 

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 and airships. 'The policy in this AC is considered applicable for airship projects; 

however, the certifying office should only use specific applicability and requirements if they are 

determined to be reasonable, applicable and relevant to the airship project. This AC applies to 

Subpart D from § 23.671 and Subpart F. This AC both consolidates existing policy documents, 

and certain ACs that cover specific paragraphs of the regulations, into a single document and 

adds new guidance. This revision has added preamble material, in italics, under the applicable 

rule and amendment level. Material in this AC is neither mandatory nor regulatory in nature and 

does not constitute a regulation. 

2. CANCELLATION or SUPERSEDING. The following AC is canceled and the policy 

statement is superseded as follows: 

a. AC 23-17A, “Systems and Equipment Guide for Certification of Part 23 Airplanes.” 

b. PS-ACE100-2002-003, “Standardization of Application of 14 CFR Part 23, § 23,1309 

Regarding Hazardous Misleading Heading Information for Attitude-Heading Reference 

Systems.” 

3. BACKGROUND. In 1968, the Federal Aviation Administration (FAA) began 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 AC covers policy available through September 30, 2003. 

Policy that became available after September 30, 2003, will be covered in future revisions to the 

AC. 

4. APPLICABILITY. This AC is applicable only to the original applicant seeking issuance of 

a Type Certificate (TC), an Amended Type Certificate (ATC), 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 AC covers policy available through September 30, 2003. 

Policy that became available after then will be covered in future revisions to the AC. 

5. PARAGRAPHS KEYED TO PART 23. Each paragraph has the applicable part 23 

amendments 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. 

.

1 
AC 23-17B 04/12/05 

a. Free Policy Memorandums, Orders and ACs. Copies of current publications of the 

following free Orders and ACs 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 ACs can be found is 

http://www.airweb.faa.gov/AC.

“PS-ACE100-2001-004, “Guidance for Reviewing Certification Plans to Address Human 

Factors for Certification of Part 23 Small Airplanes, August 29, 2002. 

PS-ACE100-2002-002, “Installation Approval of Multi-Function Displays Using the 

AML STC Process; Policy Statement,” May 21, 2003. 

PS-ACE100-2002-004, “Diesel Engine Installation,” May 22, 2003. 

FAA Order 8110.4B, “Type Certification,” April 24, 2000. 

FAA Order 8100.5A, “Aircraft Certification Service Mission, Responsibilities, 

Relationships, and Programs,” September 30, 2003. 

FAA Order 8110.42A, “Parts Manufacturer Approval Procedures,” March 31, 1999. 

AC 20-30B, “Aircraft Position Light and Anticollision Light Installations,” July 20, 

1981. 

AC 20-36S, “Index of Articles (Materials, Parts, Processes and Appliances) Certified 

Under the Technical Standard Order System,” October 1, 1993. 

AC 20-41A, “Substitute Technical Standard Order (TSO) Aircraft Equipment,” April 5, 

1977. 

AC 20-42C, “Hand Fire Extinguishers for Use in Aircraft”, March 7, 1984. 

AC 20-67B, “Airborne VHF Communications Equipment Installations,” January 16, 

1986. 

AC 20-112, “Airworthiness and Operational Approval of Airborne Systems to be Used in 

Lieu of a Ground Proximity Warning System(s) (GPWS),” February 19, 1981. 

AC 20-115B, “Radio Technical Commission for Aeronautic, Inc., Document RTCA/DO178B,” January 11, 1993. 

AC 20-118A, “Emergency Evacuation Demonstration,” March 9, 1987. 

AC 20-121A, “Airworthiness Approval of Loran-C Navigation Systems for Use in the 

U.S. National & Airspace System (NAS) and Alaska,” August 24, 1988.

2 
04/12/05 AC 23-17B 

AC 20-124, “Water Ingestion Testing for Turbine Powered Airplanes,” September 30, 

1985. 

AC 20-128A, “Design Considerations for Minimizing Hazards Caused by Uncontained 

Turbine Engine and Auxiliary Power Unit Rotor Failure,” March 25, 1997. 

AC 20-131A, “Airworthiness Approval of Traffic Alert and Collision Avoidance 

Systems (TCAS II) and Mode S Transponders,” March 29, 1993. 

AC 20-136, “Protection of Aircraft Electrical/Electronic Systems Against the Indirect 

Effects of Lightning,” March 5, 1990. 

AC 20-138A, “Airworthiness Approval of Global Navigation Satellite System (GNSS) 

Equipment,” December 22, 2003. 

AC 20-146, “Methodology for Dynamic Seat Certification by Analysis for Use in Part 

23, 25, 27, and 29 Airplanes and Rotorcraft,” May 19, 2003. 

AC 21-16D, “RTCA Document DO-160D,” July 21, 1998. 

AC 21-25A, “Approval of Modified Seating Systems Initially Approved Under a 

Technical Standard Order,” June 3, 1997. 

AC 21-34, “Shoulder Harness-Safety Belt Installations,” June 4, 1993. 

AC 23-2, “Flammability Tests,” August 20, 1984. 

AC 23-18, “Installation of Terrain Awareness and Warning System (TAWS) Approved 

for Part 23 Airplanes,” June 14, 2000. 

AC 23-19, “Airframe Guide for Certification of Part 23 Airplanes,” January 27, 2003. 

AC 23.143-1, “Ice Contaminated Tailplane (ICTS) Stall,” December 20, 2001. 

AC 23.562-1, “Dynamic Testing of Part 23 Airplane Seat/Restraint Systems and 

Occupant Protection,” June 22, 1989. 

AC 23.1309-1C, “Equipment, Systems, and Installations in Part 23 Airplanes,” March 12, 

1999. 

AC 23.1311-1A, “Installation of Electronic Displays in Part 23 Airplanes,” March 12, 

1999. 

AC 23.1419-2C, “Certification of Part 23 Airplanes for Flight in Icing Conditions,” July 

21, 2004. 

3 
AC 23-17B 04/12/05 

AC 25-11, “Transport Category Airplane Electronic Display Systems,” July 16, 1987. 

AC 90-79, “Recommended Practices and Procedures for the Use of Electronic LongRange Navigation,” July 14, 1980. 

AC 120-31A, “Operational and Airworthiness Approval of Airborne Omega Radio 

Navigation Systems as a Means of Updating Self-Contained Navigation Systems,” 

April 21, 1977. 

AC 120-37, “Operational and Airworthiness Approval of Airborne Omega Radio 

Navigation Systems as a Sole Means of Overwater Long Range Navigation,” October 10, 

1978. 

AC 121-13, and Change 2, “Self-Contained Navigation Systems (Long Range),” 

December 21, 1970. 

Copies of current publications of the following “for sale” ACs 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),” 

May 21, 1986. 

AC 20-101C, “Airworthiness Approval of Omega/VLF Navigation Systems for Use in the 

U.S. National Airspace System (NAS) and Alaska,” September 12, 1988. 

AC 21.303-2H, “Parts Manufacturer Approvals – 1992 (Microfiche),” October 13, 1992 

(Microfiche).

AC 23-16A, “Powerplant Guide for Certification of Part 23 Airplanes and Airships,” 

February 23, 2004. 

AC 23-8B, “Flight Test Guide for Certification of Part 23 Airplanes,” August 14, 2003. 

AC 43.13-1B, Change 1, “Acceptable Methods, Techniques, and Practices⎯Aircraft 

Inspection and Repair,” September 27, 2001. 

AC 43.13-2A, “Acceptable Methods, Techniques, and Practices⎯Aircraft Alterations,” 

January 1, 1977.P-8110-2, Change 2, “Airship Design Criteria,” February 6, 1995 

NOTE 1: Republishing these documents, as a part of this AC, was not considered to be the best 

utilization of FAA resources. 

b. Industry Documents

4 
04/12/05 AC 23-17B 

(1) To obtain a copy of the TSOs, write to the U.S. Department of Transportation, 

Subsequent Distribution Office, Ardmore East Business Center, 3341 Q 75th Avenue, Landover, 

MD 20785, or from the internet at http://www.faa.gov/certification/aircraft/. 

TSO-C9c, “Automatic Pilots,” September 15, 1960. 

TSO-C62d, “Tires,” September 7, 1990. 

TSO-C22g, “Safety Belts,” March 5, 1993. 

TSO-C26c, “Aircraft Wheels and Wheel-Brake Assemblies, with Addendum I,” 

May 18, 1984. 

TSO-C39b, “Aircraft Seats and Berths,” April 17, 1987. 

TSO-C55, “Fuel and Oil Quantity Instruments (Reciprocating Engine Aircraft),” 

April 1, 1959. 

TSO-C114, “Torso Restraint Systems,” March 27, 1987. 

TSO-C151b, “Terrain Awareness and Warning System,” December 17, 2002. 

(2) The RTCA documents listed below are available from RTCA, Inc., Suite 805, 1828 

L Street, NW, Washington, DC 20036: 

RTCA/DO-160D, “Environmental Conditions and Test Procedures for Airborne 

Equipment,” July 29, 1997. 

RTCA/DO-178B, “Software Considerations in Airborne Systems and Equipment 

Certification,” December 1, 1992. 

(3) The documents listed below are available from the Society of Automotive 

Engineers (SAE), Inc., 400 Commonwealth Drive, Warrendale, PA 15096: 

SAE Aerospace Recommended Practice (ARP), ARP 597C, “Wheels and Brakes, 

Supplementary Criteria for Design Endurance⎯Civil Transport Aircraft,” April 24, 

1991. 

SAE ARP 813A, “Maintainability Recommendations for Aircraft Wheel and Brake 

Design,” February 1, 1995. 

SAE ARP 1619, “Replacement and Modified Brakes and Wheels,” March 1, 2001. 

SAE Aerospace Information Report (AIR) 1064, “Brake Dynamics,” January 1, 

1997. 

5 
AC 23-17B 04/12/05 

SAE Aerospace Standard AS1145, “Aircraft Brake Temperature Monitor System

(Btms),” February 1, 1998. 

SAE J384, “Motor Vehicle Seat Belt Anchorages-Test Procedure,” June 9, 1994. 

SAE Recommended Practice, June 2004 SAE Handbook, Volume 2, 

pages 33.08-33.09. 

SAE ARP 5412, “Aircraft Lightning Environment and Related Test Waveforms,” 

November 1, 1999. 

SAE ARP 5413, “Certification of Aircraft Electrical/Electronic Systems for the 

Indirect Effects of Lightning,” (it is being incorporated into an AC), November 1, 

1999. 

SAE ARP 5414, “Aircraft Lightning Zoning,” December 1, 1999. 

SAE ARP 5415, “Users Manual for Certification of Aircraft Electrical/Electronic 

Systems for the Indirect Effects of Lightning,” May 14, 2002. 

SAE ARP 5475, “Abuse Load Testing for In-Seat Deployable-Video Systems,” 

March 1, 2001. 

SAE ARP 5577, “Aircraft Lightning Direct Effects Certification,” September 30, 

2002. 

(4) The Underwriters Laboratories (UL), Inc., document listed below can be obtained 

from Global Engineering Documents, 15 Inverness Way East, Englewood, CO 80112: 

UL 1418, “Cathode Ray Tubes,” September 6, 1996. 

(5) General Aviation Manufacturers Association (GAMA). GAMA publication can be 

purchased from the GAMA web site, http://www.gama.aero/pub: 

GAMA publication No. 10, “Recommended Practices and Guidelines for Part 23 

Cockpit/Flight Deck Design,” dated September 2000. 

c. National Aeronautics and Space Administration (NASA) document 

American Society for Testing and Materials (ASTM) Manual 36, “Safe Use of Oxygen 

and Oxygen Systems, Guidelines for Oxygen System Design, Materials Selection, 

Operations, Storage, and Transportation,” dated 2000, may be purchased from the 

ASTM, 100 Barr Harbor Drive, Post Office Box C700, West Conshohocken, PA 19428-

2959. 

6 
04/12/05 AC 23-17B 

d. European Aviation Safety Agency (EASA) Documents 

EASA Certification Specifications (CS-23) and guidance materials, Acceptable Means of 

Compliance, (AMCs), are available from the EASA website at 

www.easa.eu.int/certspecs_en.html

7. RELATED 14 CFR, PART 23 REGULATIONS 

23.75 Landing distance 

23.141 Flight Characteristics- General 

23.143 Controllability and Maneuverability- General 

23.145 Longitudinal control 

23.201 Wings level stall 

23.231 Ground and Water Handling Characteristics- Longitudinal stability 

 and control 

23.233 Directional stability and control 

23.253 High speed characteristics 

23.303 Factor of safety 

23.305 Strength and deformation 

23.307 Proof of structure 

23.335 Design airspeeds 

23.345 High lift devices 

23.351 Yawing conditions 

23.365 Pressurized cabin loads 

23.395 Control system loads 

23.397 Limit control forces and torques 

23.405 Secondary control system

23.473 Ground load conditions and assumptions 

23.477 Landing gear arrangement 

23.479 Level landing conditions 

23.481 Tail down landing conditions 

23.483 One-wheel landing conditions 

23.485 Side load conditions 

23.493 Braked roll conditions 

23.505 Supplementary conditions for ski-planes 

23.521 Water load conditions 

23.561 Emergency Landing Conditions- General 

23.562 Emergency landing dynamic conditions 

23.601 Design and Construction- General 

23.603 Materials and workmanship 

23.613 Material strength properties and design values 

23.655 Control Surfaces- Installation 

23.955 Fuel flow 

23.959 Unusable fuel supply 

23.961 Fuel system hot weather operation 

7 
AC 23-17B 04/12/05 

23.991 Fuel pumps 

23.993 Fuel system lines and fittings 

23.995 Fuel valves and controls 

23.997 Fuel strainer or filter 

23.1019 Oil strainer or filter 

23.1121 Exhaust System- General 

23.1123 Exhaust system

23.1182 Nacelle areas behind firewalls 

23.1183 Lines, fittings, and components 

23.1191 Firewalls 

23.1203 Fire detector system

23.1505 Airspeed limitations 

23.1525 Kinds of operation 

23.1527 Maximum operating altitude 

23.1529 Instructions for Continued Airworthiness 

23.1541 Markings and Placards- General 

23.1543 Instrument markings: General 

23.1545 Airspeed indicator 

23.1547 Magnetic direction indicator 

23.1549 Powerplant and auxiliary power unit instruments 

23,1551 Oil quantity indicator 

23.1553 Fuel quantity indicator 

23.1555 Control markings 

23.1557 Miscellaneous markings and placards 

623.1559 Operating limitations placard 

23.1581 Airplane Flight Manual and Approved Manual material- General 

23.1583 Operating limitations 

23.1585 Operating procedures 

Appendix F to Part 23 Test Procedure (Flammability) 

8 
04/12/05 AC 23-17B 

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 September 30, 2003. EASA AMC 23.671 is acceptable for 

FAA certification. 

This rule was adopted on February 1, 1965 as a recodification of Civil Air Regulations 

(CAR), CAR 3.335. 

9 
AC 23-17B 04/12/05 

23.672 Stability augmentation and automatic and power-operated systems 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

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. 

A proposed revision to NPRM 90-18 explains this amendment as follows: “This 

proposal would provide criteria for approval of those stability augmentation, 

automatic, and power-operated systems whose performance is essential to flight 

safety. The proposed § 23.672 is similar to § 5.672 and, as part 25, the warning 

system requirement relating to control system activation is not intended to preclude 

installing tactile warning devices, such as control system shakers activated 

independently for other purposes.” 

NOTE 2: Section 5.672 in above quote is incorrect. It should be § 25.672. 

This requirement would not apply to a simple downspring or a bobweight stability 

device as affirmed by Final Rule, Docket 26269.

10 
04/12/05 AC 23-17B 

23.673 Primary flight controls 

No policy available as of September 30, 2003. 

The corresponding rule in CAR 3 is CAR 3.336. 

11 
AC 23-17B 04/12/05 

23.675 Stops 

No policy available as of September 30, 2003. 

The corresponding rule in CAR 3 is CAR 3.340.

12 
04/12/05 AC 23-17B 

23.677 Trim systems 

The corresponding rule in CAR 3 is CAR 3.337.

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.16.

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. 

This amendment is clarified by Final Rule, Docket 8083 as follows: “The notice 

proposed to amend § 23.677 to make it clear that operation of the trim system may 

not be dependent upon the primary control system and that failures of the primary 

control system may not prevent safe flight and landing. Comments were received 

objecting to the proposal on the ground that it exceeded the part 25 requirements for 

transport category airplanes, and that it would improperly and undesirably change 

trim response and make compliance tests extremely hazardous. The FAA does not 

agree. The trim requirement is merely one of a number of control system 

requirements that must be considered as a total requirement. There is, therefore, no 

comparison between individual provisions in parts 23 and 25. Service experience 

with existing airplanes does not indicate that the proposal will improperly and 

undesirably change trim response or that compliance testing will be hazardous. 

Furthermore, trim response equal to primary flight control response will not 

necessarily be needed to comply with the proposal. All that is required is that there 

be adequate control for safe flight and landing.” 

13 
AC 23-17B 04/12/05 

Amendment 23-34 and Subsequent 

Section 23.677(d) was added by Amendment 23-34 was only applicable to commuter 

category airplanes. 

Probable powered trim runaways should be demonstrated for all part 23 airplanes so 

equipped. See AC 23-8B, “Flight Test Guide for Certification of part 23 Airplanes,” 

for the procedure. 

Even if trim runaways have been determined to be unlikely 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 the 

airplane has no unsafe features. The FAA has accepted demonstration of controlrestrained 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 unlikely. However, the FAA has determined this procedure is 

not acceptable in itself for failure conditions shown to be less than unlikely. To allow 

expansion of the 0 to 2g envelope, as specified in AC 23-8B, the FAA suggests a test 

procedure that incorporates both control restrained and unrestrained errors. 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 (that is, feet 

on the floor), restrained runaways for rudder trim are not considered acceptable. (See 

Table 1). 

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)

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). 

2. Trim systems with a monitor/limiter will be tested at a magnitude just below that required for 

monitor/limiter trip. 

3. NTE is Not to Exceed. 

14 
04/12/05 AC 23-17B 

Amendment 23-42 and Subsequent 

This section was changed by Docket 25811, Amendment 23-42, as follows: “This 

proposal would extend the current requirements of Sec. 23.677 for powered trim 

system runaways to all categories of part 23 airplanes.” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21, Amendment 23-49, revised § 23.677(a) as 

follows: 

“Proposed revised Sec. 23.677(a) would clarify the need to mark the lateral and 

directional trim indicators with the neutral trim position. Since trim indicators on 

most airplanes are currently marked with the neutral position of the trimming device, 

this proposal would standardize the cockpit markings for all airplanes. 

Revised paragraph (a) would also add a requirement for the pitch trim indicator to 

be marked with the proper pitch trim range for the takeoff of the airplane. Some 

takeoff accidents, including some involving fatalities, have occurred because the 

pitch trim was not set to the proper range needed for the airplane takeoff. Because of 

this accident experience, most of the current airplane manufacturers mark the pitch 

trim indicator with the pitch trim range for takeoff. Therefore, the proposed marking 

requirement would not have a significant impact on future airplane designs and 

would ensure that the markings needed for a safe takeoff are provided for the pilots 

use.”

15 
AC 23-17B 04/12/05 

23.679 Control system locks 

The corresponding rule in CAR 3 is CAR 3.341. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.17.

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 obvious warning to 

the pilot when the lock is engaged. Several accidents have occurred because the pilot 

did not remove the control system lock before 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, and 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 sothe 

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 (for 

example, throttle) can be removed independently of the other, each lock should 

independently meet the criteria of paragraph (c) above. 

16 
04/12/05 AC 23-17B 

Amendment 23-45 and Subsequent 

This amendment, changed by Final Rule, Docket 26269, states as follows: “The FAA 

is aware that an automatically released control lock system would be costly. The 

proposal did not mandate the installation of an automatic system, but would add an 

optional provision that would show the acceptance of such systems. 

The JAA stated its assumption that the proposed requirement would not be applicable 

to external locks. Based on the comments received, the FAA has re-examined the 

proposal. Since the proposal would have eliminated the current Section 23.679(a), 

external systems that use the red warning ribbons as a means of warning the pilot 

that the locks are in place would no longer be acceptable. The FAA has determined 

that there is a need to retain the provision of current Section 23.679(a), so that 

presently used locks and their warning systems remain acceptable. The added 

provision of Section 23.679(a)(2) will make it clear that systems that automatically 

disengage the locks are also acceptable but not mandatory. 

The proposal to limit the operation of the airplane when the locks are engaged is 

being restated since control locks and their warnings can be overlooked and 

automatic disengage systems will fail. The FAA believes an additional safeguard is 

required. By requiring a system that will ensure that airplane operation is limited, 

the pilot will receive a pre-takeoff warning and thus a hazardous takeoff will not be 

attempted. 

In summary, the FAA has considered the comments and has revised the proposed rule 

language by retaining the current provisions of Section 23.679(a) and Section 

23.679(a)(1), and by adding the provision for accepting automatically disengaged 

locking systems as an option. The language in proposed Section 23.679(a)(2) to 

require the control surfaces to be locked so the pilot receives an unmistakable 

warning at the start of the takeoff if the locks have not been removed is retained as 

Section 23.679(b). The unmistakable warning required by this paragraph may be a 

tactile warning that the pilot receives by the feel of the controls. Finally, proposed 

Section 23.679(b) is retained as paragraph (c).”

17 
AC 23-17B 04/12/05 

23.681 Limit load static tests 

No policy available as of September 30, 2003. 

The corresponding rule in CAR 3 is CAR 3.342. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.18. 

18 
04/12/05 AC 23-17B 

23.683 Operation tests 

The corresponding rule in CAR 3 is CAR 3.343. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.19. 

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 

A proposed revision to NPRM 67-14 explains this amendment as follows: 

“This proposal would preserve consistency with Sec. 23.397 which provides that the 

airloads on the surfaces need not exceed those that would result from application of 

the forces in Sec. 23.397 (b).” 

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, 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. Second, 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 § 3.343 is as follows:

19 
AC 23-17B 04/12/05 

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 

pilots 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: 

(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 pilots 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 torques 

specified by the regulations may be met. 

ALTERNATE MEANS OF COMPLIANCE 

Applicants and FAA Aircraft Certification Offices (ACOs) involved with certification 

of small airplanes should follow this policy. Applicants should expect the ACO 

would consider this information when making findings of compliance. 

20 
04/12/05 AC 23-17B 

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 assure 

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 method also can have the added 

advantage of showing compliance with several regulations in a single test series. 

Regardless of the amount of travel of a control surface when tested as, 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 assure 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 clearly 

specify 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: 

1. 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, 

"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 as follows: 

(1) The direction of the test loads produces the most severe loading in the 

control system; and 

21 
AC 23-17B 04/12/05 

(2) Each fitting, pulley, and bracket used in attaching the system to the main 

structure is included." 

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 assure 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, sometimes 

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 

case-by-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 show 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 show 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. 

22 
04/12/05 AC 23-17B 

(2) 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 may 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 show compliance with both 

§ 23.305, paragraph (a), and § 23.683. It also shows compliance with 

§ 23.681, paragraph (a). These tests may be conducted as follows: 

Except where otherwise mentioned, 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 pilots' 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 

cable tension, decreased cable tension, and control/fixed surface 

proximity for each control system as appropriate. 

23 
AC 23-17B 04/12/05 

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)). For § 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 airplanes 

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 because of 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 found 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). 

(iv) Load the pilots control until the control surface is just off the support. 

24 
04/12/05 AC 23-17B 

(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 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 paragraphs d 

and e above. 

(ii) Operate the unloaded control system from stop to stop. 

(iii) No signs of interference (contact) may be present. 

25 
AC 23-17B 04/12/05 

(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 3: 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 for type design. A sufficiently representative 

mockup of remaining control system components must be used to assure the full 

lengths 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 show 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 show 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 found 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 pilots 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. 

(vi) The minimum loaded control surface travel must have adequate flight 

characteristics as specified in § 23.141.

26 
04/12/05 AC 23-17B 

(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 4: The tests described in section (3) above are normally performed using a 

complete airplane. As a minimum, they must be completed using an airframe/control 

system that represents the final product from the cockpit controls to the control 

surface. 

Regardless of the amount of travel of a control surface when tested , 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 assure compliance. 

EASA AMC 23.683 is acceptable for FAA certification. 

27 
AC 23-17B 04/12/05 

23.685 Control system details

No policy available as of September 30, 2003. 

The corresponding rule in CAR 3 is CAR 3.344. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.20. 

28 
04/12/05 AC 23-17B 

23.687 Spring devices

This rule was adopted on February 1, 1965 as a recodification of CAR 3.347. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.21. 

Compliance may be shown by flight tests with the spring disconnected to show that 

failure of the spring will not cause flutter or unsafe flight characteristics, or by 

performing a reliability analysis. 

The flight test is preferred since this approach clearly addresses the safety issue and is 

historically a minimal test burden. 

An applicant who chooses reliability analysis 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 showing the reliability 

of the spring device. The reliability assessment should consider, but not necessarily 

be limited to, fatigue failures, failures because of corrosive environments, and any inservice changes in the spring characteristics, 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 clear in that the applicant must show 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. Also 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). 

29 
AC 23-17B 04/12/05 

23.689 Cable systems 

The corresponding rule in CAR 3 is CAR 3.345. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.22.

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. They are 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. 

A proposed revision to NPRM 67-14 revised this section by Amendment 23-7 as 

follows: “Pulley specifications (Sec. 23.689). Section 23.689(b) was amended by 

striking out the words "as specified in the pulley specifications" after the words "is 

used. Explanation. Correspondence of pulley and cable is prescribed in the first 

sentence of Sec. 23.689(b). Correspondence may be verified in several ways. The 

reference to pulley specifications adds nothing.” 

30 
04/12/05 AC 23-17B 

23.691 Artificial stall barrier system 

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Amendment 23-49 and Subsequent

A proposed revision to NPRM 94-21 states as follows: “This proposed new section 

would provide standards for stall barrier systems if a stall barrier is necessary to 

show compliance with Sec. 23.201(c). 

The requirements of Sec. 23.201(c) provide criteria for the in-flight demonstration of 

wings level stall. The requirements also specify the means of identifying when a stall 

has occurred. Amendment No. 23-45 (58 FR 42136, August 6, 1993) revised Sec. 

23.201(c) by adding the activation of an artificial stall barrier as an acceptable 

means of identifying when a stall has occurred. 

As the technology of airplane designs improved and engines with increased power 

became available, airplanes were developed that did not meet the older wings level 

stall requirement of Sec. 23.201. Consequently, these airplanes were equipped with 

an artificial stall barrier that moved the airplane elevator controls and caused a nose 

down pitching motion similar to the pitching motion of airplanes that meet the wings 

level stall requirement of Sec. 23.201. The manufacturer selected the airspeed where 

this pitching motion occurred and flight testing established compliance with the other 

flight regulations at airspeeds above the speed selected for the push. These stall 

barrier systems are commonly called "stick pushers." Such systems have been 

accepted for compliance with Sec. 23.201 under the equivalent safety provisions of 

Sec. 21.21(b)(1), since they provide a pitch motion that is equivalent to that 

experienced during stalls of airplanes that meet the stall requirements of Sec. 23.201. 

Appropriate compliance with other applicable requirements of part 23 has been 

established by other design characteristics of the stall barrier system. 

The provisions of the proposed new section are based on system design 

characteristics necessary to ensure the safe operation of previously approved stall 

barrier systems. The proposed section also requires such systems to include 

provisions to prevent unwanted activation of the stall barrier systems. This is 

necessary to ensure that such systems do not cause downward pitching motions at 

higher airspeeds when such pitching could be unsafe. 

The proposed sections would basically codify those provisions that have been found 

necessary for approving stick pusher systems under the equivalent safety 

requirements of Sec. 21.21(b)(1). Therefore, in effect, no new requirements would be 

added by this proposed amendment. 

31 
AC 23-17B 04/12/05 

The proposed new section would be applicable only to airplanes with flight 

characteristics that need an artificial stall barrier system to ensure safe operation of 

that airplane. Including provision for the installation of an optional stick pusher 

system would relieve the manufacturer of the financial burden that would be needed 

to redesign the airplane so that it would meet the wings level stall requirements.” 

The proposed rule was changed by Final Rule, Docket 27806 as follows: “In the 

course of the FAAs review, however, the FAA noted that the word "necessary" in the 

introductory paragraph of Sec. 23.691 should be changed to "used," to make it clear 

that the equipment requirements of this section are applicable if a stick pusher system 

is used in the airplane to show compliance with Sec. 23.201(c).” 

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 (that is, a stick pusher) while a stall 

warning is a device that alerts a pilot of an impending stall (that is, a stick shaker). 

Of course, the 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). 

For § 23.201, a stick shaker is a "stall avoidance" device. In 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 Airplane Flight Manual 

(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 unlikely. The FAA does not consider the chance 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.) 

To develop 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. 

32 
04/12/05 AC 23-17B 

Stall can be identified by stick shaker/pusher operation, uncontrollable downward 

pitching, or the elevator control reaching the stop (see AC 23-8B) -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 unlikely. 

33 
AC 23-17B 04/12/05 

23.693 Joints

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.346. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.23. 

34 
04/12/05 AC 23-17B 

23.697 Wing flap controls

The corresponding rule in CAR 3 is CAR 3.338. 

There is no corresponding rule in the Airship Design Criteria. 

Amendment 23-49 and Subsequent 

A revision to NPRM 94-21 proposed § 23.867(c) as follows: “This revision is 

needed to ensure that the flap settings, which establish the safe operation of the 

airplane, can be positively selected.” 

35 
AC 23-17B 04/12/05 

23.699 Wing flap position indicator

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.338. 

There is no corresponding rule in the Airship Design Criteria. 

36 
04/12/05 AC 23-17B 

23.701 Flap interconnection 

The corresponding rule in CAR 3 is CAR 3.339. 

There is no corresponding rule in the Airship Design Criteria.

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 allowed 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 shown as safe with one hand on 

the control wheel/stick (other hand needed to re-trim, pull circuit breaker, operate flap 

control,.). 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 showing 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

A proposed revision to NPRM 89-5 explains this amendment as follows: “This 

proposal would update the regulations to include provisions for airplanes with a flap 

configuration other than one flap on each wing. It would also address the failure of 

any single element in the flap control system and allow for an alternate equivalent 

means to the mechanical interconnection required by the present rule. Airplanes are 

currently being manufactured with two flaps on each side of the airplane and are 

being designed with flaps on canards and tandem wings. On an airplane with four 

flaps, there is a possibility that only one flap may be asymmetric with respect to the 

other three and this issue needs to be addressed in the airworthiness standards, as 

proposed in paragraphs (a) and (b).” 

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 the event of an unsafe condition has been 

shown to be unlikely. 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.

37 
AC 23-17B 04/12/05 

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 event of an unsafe condition has been 

shown to be unlikely; or 

(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 the new § 23.701(b), 

particularly the parenthetical portion of that paragraph, could be improperly 

interpreted and applied. It could be possible that this misinterpretation could result in 

the use of differing terminology (that is, "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 (ELOS) 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. 

38 
04/12/05 AC 23-17B 

b. Mechanical Interconnection Requirement of § 23.70l(a)(l) 

This requirement is to assure 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 show 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. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 revised this section as follows: “Section 23.701 

(a)(1) and (a)(2) would be revised to clarify the requirements for flap systems 

installed on part 23 airplanes. Following the revision of Sec. 23.701, as adopted by 

Amendment No. 23-42 (56 FR 353, January 3, 1991), the FAA discovered that the 

new requirements could be interpreted in a way that was not intended and that this 

interpretation could result in approval of airplanes with unsafe flight characteristics 

in the event of flap failure. To clarify the intent of the requirements, the FAA issued 

on March 14, 1991, a policy letter to all aircraft certification offices that provided 

guidance for the correct application of the requirements. 

Since then, the FAA has reexamined the requirements and determined that Sec. 

23.701(a)(1) and (a)(2) need to be revised to ensure that a failure of the flap system 

would not create an asymmetric flap configuration that could result in an unsafe 

flight condition. Therefore, Sec. 23.701 (a)(1) and (a)(2) would be revised to clarify 

that one of the following would apply: 

(1) The moveable flap surfaces must be synchronized by a mechanical 

interconnection or by an approved equivalent means that is independent of the 

flap drive system. 

(2) The wing flap system must be designed so that any failures of the flap system 

that would result in an unsafe flight characteristic of the airplane, such as flap 

asymmetry, are extremely improbable. 

These revisions would ensure that a failure of the flap drive systems will not result in 

a flap asymmetry configuration.” 

39 
AC 23-17B 04/12/05 

ACCEPTABLE MEANS OF COMPLIANCE

An acceptable means of compliance with the airworthiness requirements for the flaps 

mechanical interconnections of § 23.701(a)(1) are described as follows: 

a. Reliability 

Reliability of the mechanical interconnections is 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 

because 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 

a hazardous flap asymmetrical operation would not occur after a jam condition, 

the 1.5 factor of safety should not be required.

40 
04/12/05 AC 23-17B 

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 (that is, 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 

on design safety assessments, environmental and atmospheric conditions, and 

software assessment. 

41 
AC 23-17B 04/12/05 

23.703 Takeoff warning system

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-49 and Subsequent 

A revision to NPRM 94-21 proposed this rule as follows: “This proposed new 

section would require a takeoff warning system on some commuter category 

airplanes. The requirement would be applicable if the flight evaluation showed that 

an unsafe takeoff condition would result if lift devices or longitudinal trim devices are 

set to any position outside the approved takeoff range. If the evaluation shows that 

no unsafe condition would result at any setting of these devices, a takeoff warning 

system would not be required. For those airplanes on which a warning system would 

not be required. For those airplanes on which a warning system must be installed, 

the proposal would provide requirements for the installation of the system.” 

42 
04/12/05

AC 23-17B 

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. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

43 
AC 23-17B 04/12/05 

23.723 Shock absorption tests 

The corresponding rules in CAR 3 are CAR 3.351 and 3.352. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.24.

Original Issue

This regulation requires shock absorption tests be performed for certification. 

Amendment 23-23 and Subsequent

This amendment allows 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 (for example, structure, wheel tire, shock absorber) should 

be included in determining the 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. 

44 
04/12/05 AC 23-17B 

23.725 Limit drop tests 

The corresponding rules in CAR 3 are CAR 3.353 and 3.354. 

There is no corresponding rule in the Airship Design Criteria.

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 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. 

45 
AC 23-17B 04/12/05 

23.726 Ground load dynamic tests

No policy available as of September 30, 2003. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

46 
04/12/05 AC 23-17B 

23.727 Reserve energy absorption drop test 

The corresponding rule in CAR 3 is CAR 3.355. 

There is no corresponding rule in the Airship Design Criteria.

Amendment 23-7 and Subsequent

Paragraph (b) in § 23.727 requires 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). 

47 
AC 23-17B 04/12/05 

23.729 Landing gear extension and retraction system 

The corresponding rules in CAR 3 are CAR 3.356, 3.357, 3.358 and 3.359. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.25.

Original Issue 

14 CFR, part 23, does not define landplanes, seaplanes and amphibians. These terms 

passed unchanged from CARs 3 and 4. As we use them, landplanes can only land on 

the ground, seaplanes (both boat and float types) can only land on water, and 

amphibians can land on both water and on the ground. 

Therefore, where a rule such as § 23.729(c) states "landplanes" it does not apply to 

seaplanes or amphibians. When the rule states "airplanes" such as § 23.729(a) it 

applies to any aircraft that incorporates that equipment. Thus, § 23.729(a) can apply 

to landplanes and amphibians, but not seaplanes since they don't have retractable land 

gear. 

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 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 suggest 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,. This logic may minimize nuisance 

warnings and may provide the equivalent safety intended by the rule. 

Because part 23 is not specific about 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 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.

48 
04/12/05 AC 23-17B 

Amendment 23-21, NPRM 75-25, changed this section to require a warning when 

any throttle is closed. The NPRM explanation states as follows: “The present 

rule requires that a manual shutoff for the aural landing gear warning device be 

installed so that reopening the throttle will reset the device. However, when an 

engine has been shut down in flight, its throttle may not be reopened before 

landing. The proposal would require the aural warning to be activated when any 

throttle is subsequently retarded to or beyond the position for a normal landing 

approach, thus requiring the warning intended by paragraph (e)(2) regardless of 

the position of any other throttle and the prior deactivation.” 

Amendment 23-45, NPRM 90-18, changed this section again. The NPRM states 

as follows: “This proposal revises § 23.729(f)(1) and (2) by changing the power 

and flap settings necessary to activate the device that warns the pilot that the 

landing gear is not fully extended and locked. The power setting necessary to 

activate the warning device is changed from when one or more "throttles are 

closed" to when one or more "throttle are closed beyond the power settings 

normally used for landing approach." The flap setting necessary to activate the 

warning device is changed from "flaps are extended to or beyond the approach 

flap position" to " flaps are extended beyond the approach flap position." 

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 for 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 

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:

49 
AC 23-17B 04/12/05 

(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 15 degrees. 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 so that Vq is 

less than VA. 

(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 intent 

of the proposed revision to NPRM 94-21 is given as follows: “This proposal would 

also add a new Sec. 23.729(g) requiring that if the landing gear bay is used as the 

location for equipment other than landing gear, the equipment must be designed and 

installed to minimize damage. On larger airplanes, such as the commuter category, a 

50 
04/12/05 AC 23-17B 

primary cause of damage to such equipment would be tire-burst. In addition, service 

history has shown that rocks, water, and slush enter the landing gear bay and cause 

damage. The equipment on any size airplane should be protected from damage by 

such external sources.” 

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 § 23.729(g) compliance. Systems that act on the landing gear, for 

example, extension and retraction systems, hydraulic systems, 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 AC 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 airplane safety 

analysis is consistent with the FAA advisory material for parts 23, 25, and 29, 

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 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 answers to the questions raised as follows: 

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)?”

51 
AC 23-17B 04/12/05 

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. 

EASA AMC 23.729(g) is acceptable for FAA certification. 

52 
04/12/05 AC 23-17B 

23.731 Wheels 

See 23.735, Brakes. 

The corresponding rule in CAR 3 is CAR 3.361. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.26.

53 
AC 23-17B 04/12/05 

23.733 Tires

The corresponding rule in CAR 3 is CAR 3.362. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.27.

Original Issue and Subsequent 

The following is a recommended test procedure for installing 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 efforts; 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 5: Above applies to a dual tire installation for 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. 

A revision to Amendment 23-17, NPRM 75-10, added subparagraph (c) with the 

following explanation: “The proposed rule would require that the selection of tires 

for installation on retractable landing gear mechanisms take into account the tire 

production tolerance and size increases that would be expected to result from service. 

The FAA believes compliance with the proposed rule could prevent accidents that 

might result from jamming of landing gear mechanisms by oversize tires.” 

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 FAA to examine the effects of tundra tires installed 

on a Piper PA-18. Second, it provides information on 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 

54 
04/12/05 AC 23-17B 

an example "compliance checklist" for the installation of such tires on light 

airplanes that have 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. 

c. Technical Standard Order (TSO)-C62d, “Aircraft Tires.” 

d. AC 43.13-1B, Change 1, “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 the possibility of problems with tundra tires be investigated. 

The NTSB stated the following: 

“Since the early 1960s, 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 STCs 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-18s 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 FAAs 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 for safety. Rate of climb and cruise speed were degraded with the 

larger tire sizes; however, the aircraft still met certification requirements. 

55 
AC 23-17B 04/12/05 

Added 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 characteristics are further aggravated when 

operators of the PA-18 remove the 2.5 degrees of washout at the wing tip, which 

is not an approved change. 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.

56 
04/12/05 AC 23-17B 

(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. 

(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: 1) reduce the airplane's directional stability and controllability 

during takeoff and landing ground rolls, 2) increase its tendency to ground loop 

during takeoff and landing ground rolls, and 3) 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. 

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.

57 
AC 23-17B 04/12/05 

(6) Changes in tires may affect the air induction system certification for 

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 either by 

TC or STC. The applicable requirements for installation of a tire on a given 

airplane should be determined based on the original certification basis 

specified in that airplane's Type Certificate Data Sheet (TCDS). Developing a 

compliance checklist, as described in item 7 below, should be accomplished 

by the applicant and the FAA engineer. 

b. An alternative certification method exists for a tire that does not have a TSOA. 

The tire design approval may be obtained concurrently with the installation 

approval for specific airplane models by 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 on the original certification basis 

specified in that airplane's TCDS. Developing a compliance checklist, as 

described in item 7, should be accomplished by the applicant and the FAA 

engineer. Before 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 

compliance checklist for a specific installation should be determined at the start of 

a project. (See Appendix 2 in this section, which is applicable to § 23.733 Tundra 

Tires.) 

58 
04/12/05 AC 23-17B 

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 NTSB 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 (CG)/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 testing the four tundra tires were 

compared to the data obtained from testing 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 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 obvious when making a main wheel only 

landing. Crosswind landings on runway 13 at Lake Hood Strip, a 2,200 feet x 

80 feet gravel runway next to Lake Hood three miles southwest of Anchorage, 

Alaska, were demonstrated for tire configurations items 1a, 1b, and 1c in 

winds from 180 degrees (from ahead and to the right of the airplane at an 

angle of 50 degrees to its flight path) at 14 knots gusting to 16 knots. The 

wind thus had a crosswind component of approximately 10.7 knots gusting to 

approximately 12.3 knots and a head wind component of approximately 9.0 

59 
AC 23-17B 04/12/05 

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. 

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 confirm 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 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 (for example, while circling to spot moose), the airplane may 

impact the ground before 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 degrees of washout at the wing tip, which is not an approved change. This 

condition will result in a rapid roll when the airplane is stalled during turning 

flight. 

4. Handling Qualities 

For any given CG/weight, the lateral and directional stability deteriorates as tire 

size is increased. 

60 
04/12/05 AC 23-17B 

5. Stall Warning 

Installation of the artificial stall warning system on the PA-18 is optional. Most 

of the PA-18s in Alaska do not have the system installed. The airplane tested did 

have the artificial stall warning system, and several test points were obtained with 

the system deactivated. The airplane as tested does not have an aerodynamic stall 

warning. 

A1-61 (and 62) 

04/12/05 AC 23-17B 

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 CG 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*

63 
AC 23-17B 04/12/05 

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

64 
04/12/05 AC 23-17B 

65 

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 
AC 23-17B 04/12/05 

66 

Subpart F—EQUIPMENT 

Landing Lights 

3.699 Landing light installation 

Subpart G—OPERATING LIMITATIONS AND INFORMATION 

3.735 General 

Limitations 

3.737 Limitations
04/12/05 AC 23-17B 

67 

23.735 Brakes 

The corresponding rule in CAR 3 is CAR 3.363. 

There is no corresponding rule in the Airship Design Criteria.

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 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
AC 23-17B 04/12/05 

68 

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 “Parts Manufacturer Approvals”⎯October 13, 1992 

(Microfiche). 

AC 23-8B “Flight Test Guide for Certification of Part 23 

Airplanes,” August 14, 2003. 

c. Technical Standard Order (TSO) 

TSO-C26c “Aircraft Wheels and Wheel-Brake Assemblies, 

with Addendum I,” May 18, 1984. 

d. Industry Documents 

SAE ARP 597 “Wheels and Brakes, Supplementary Criteria Design 

for Endurance-Civil Transport Aircraft,” April 24, 

1991. 

SAE ARP 813 “Maintainability Recommendations for Aircraft 

Wheels and Brake Design,” February 1, 1995. 

SAE AIR 1064 “Brake Dynamics,” January 1, 1997. 

SAE AS 1145 “Aircraft Brake Temperature Monitor System 

(Btms),” February 1, 1998. 
04/12/05 AC 23-17B 

69 

SAE ARP 1619 “Replacement and Modified Brakes and 

Wheels,” March 1, 2001.

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 on issues that have been identified by 

industry under ARP 1619, which, in part, concerns the variance in compliance provisions 

associated with original equipment manufacturers (OEM) and non-OEM applicants. The 

guidelines will include: 1) a description of replacement or modified wheel, 2) 

wheel/brake parts or assembly changes, 3) a description and examples of associated 

"major" and "minor" changes, and 4) 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. This will result in an equivalent level of certified type performance and 

safety to that shown by either the originally approved parts or assemblies. The 

change may be approved under a 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 showing 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 TSOA, TC, or an STC with the submission of any design data covered 

by the licensing agreement, as determined by the FAA. 

(2) Identicalness 

Applicant should provide evidence the parts he produces will be identical to the 

corresponding parts of an approved 14 CFR type design, TSO, or PMA. 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.
AC 23-17B 04/12/05 

70 

(3) Airworthiness Requirements (Tests and Analyses) 

Applicant should provide evidence—in the form of drawings, test reports, 

computations, and other substantiating data—showing 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: 

(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 

 AFM. 

(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-Brake 

 Assemblies, with Addendum 1” 

NOTE 6: A description of the certification basis in which an airplane was approved can be 

obtained from the FAA.
04/12/05 AC 23-17B 

71 

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 shown 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 showing compliance to methods identified under 

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 as 

follows: 

(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 greatly 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. 
AC 23-17B 04/12/05 

72 

(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: 1) one that would have no noticeable affect on either the 

performance of the original TSO assembly or 2) 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 the change can be made without adversely 
04/12/05 AC 23-17B 

73 

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.

a. Substantiation Plan 

A proposed substantiation plan may be presented to the FAA for approval that 

identifies the applicants requested change and intended approach in substantiating the 

change in accordance with the methods addressed under this section. The plan should 

include: 

(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" and the 

basis for the classification about 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 used. 

(8) The quality management and quality assurance system under which either the part 

or assembly, or both, will be produced. 

b. Substantiation Requirements 

The recommended substantiation requirements for replacement and modified wheels, 

wheel/brake parts or assemblies are based on changes for which approval is sought, 

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 AFM performance data, then 

the applicant should provide the applicable performance data in an FAA approved 

AFM or Airplane Flight Manual Supplement (AFMS). Depending on 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 by 

compliance with the following substantiation requirements:
AC 23-17B 04/12/05 

74 

(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: 

1) they are proposed by the original wheel and brake manufacturer who holds 

the TSO authorization or 2) 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 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 by 3(d)(1), titled 

“Major Design Changes,” unless it can be shown the change cannot affect: 1) 

the airplane stopping performance, 2) brake energy absorption characteristics, 

or 3) 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, 

and vibration control, should also be considered for the major/minor 

determination. The original manufacturer of the wheel or wheel/brake 

assembly who holds the TSO authorization may possess sufficient data to 

show that such changes could be considered minor (for example, 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 enough data to show that performance would not be 

affected. In this case, the major/minor status would be determined by 

applicable dynamometer tests by 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: 1) a controlled test at the maximum certified kinetic 

energy capacity of the original brake assemblies from TSOA, or 2) 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 
04/12/05 AC 23-17B 

75 

validate replacement rotors/stators proposed by an applicant other than the 

original TSO holder:

1. Use of new stator or rotor parts in the replacement manufacturers brakes 

for each dynamometer test will be required to minimize test configuration 

variables. If rebuilt or in-service components other than these fail during 

testing, it should be realized the results may be questionable. Suspect tests 

would be inspected 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 KEs approved under TSOA for the original 

manufacturer are proprietary data. Therefore, a PMA applicant that wants 

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 manufacturers brake. For each succeeding run, the KE will be 

increased by at least five 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 functional landings for anti-skid equipped airplanes and 

a minimum of three functional landings for non-anti-skid equipped 

airplanes, as, are needed.

(d) Worn Brakes [optional] 

While there are no provisions under part 23 to require evaluating brake 

performance using worn brakes, there have been rejected takeoff accidents in 

which the brakes on subject airplanes were at or near their worn state of 

energy absorption capability and stopping capability. Therefore, as an 

optional test to a replacement brake performance evaluation (when there is 

question about variances in worn brake performance), it is recommended that 

such an assessment on the dynamometer be undertaken to support compliance 

with 
AC 23-17B 04/12/05 

76 

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 (for example, at least 90 percent worn). The tests, used to verify 

the safety of a replacement brake system and to determine the maximum 

energy absorption capability of brakes in their fully worn state, should be 

substantiated as being representative of airplane and runway conditions. 

(2) Modified Wheel, Wheel/Brake Parts, or Assemblies 

(a) Modified Brake Design 

This laboratory and airplane test requirement applies to adding 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: 

1. For improved performance

(aa) Applicable dynamometer tests under TSO-C26c. 

(bb) Instrument flight tests to include six takeoffs and six 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 that may be applicable 

to the test.

(cc) Additional tests may be necessary for each airplane configuration 

change (for example, takeoff flaps, landing flaps, nose wheel 

brakes, anti-skid devices inoperative, and deactivation of 

wheel/brakes). 

(dd) Brake system response evaluation. 

(ee) Parking brake adequacy. Tires are allowed to skid during 

maximum power engine checks. 

(ff) Alternate braking system stops. 
04/12/05 AC 23-17B 

77 

(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. 

(jj) At least two 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 7: Improved performance implies an increase in the friction coefficient (mu) versus 

energy level for the desired operation(s) and may be requested for landing, RTOs, 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 

AFM 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 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 8: Equivalent performance implies that enough data will be obtained to verify the 

performance level for the change is equal to or better than the existing performance levels. 

The change may be to change the CG 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). 
AC 23-17B 04/12/05 

78 

NOTE 9: Extended performance implies 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). 

(b) Modified Anti-skid System 

This airplane test requirement applies to adding a new anti-skid system or 

changes to an existing anti-skid system that may affect airplane performance 

(for example, 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. If 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 adding 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 the fuse plugs 

would 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 acceptable for this test. 

NOTE 10: Wheel fuse plug integrity should be confirmed during braking tests where the 

energy level simulates the maximum landing energy. It should be shown 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 that 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). 
04/12/05 AC 23-17B 

79 

(dd) Conduct accelerate-stop test at maximum landing energy, keeping 

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).

(ff) Park in an area to minimize wind effects until it is assured 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, a landing and quick turnaround may be 

performed. Caution should be exercised to prevent jeopardizing the 

safety of the flight crew and airplane if the wheel plugs release 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, keeping 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. Go 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 assured that fuse 

plug temperatures have peaked and that no plugs have released. Fuse 

plug protection of wheels and tires should show the fuse plugs would 

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 to 
AC 23-17B 04/12/05 

80 

certify 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.”

(e) Other Substantial Airplane Tests 

Depending on the extent of wheel, wheel/brake part, or assembly 

modifications that may be involved, there will be several airplane tests that 

should be considered as well as 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 Kinetic Energy (KE) Absorption Tests 

Verify the brake KE absorption test determined by the laboratory test 

meets the TSO requirements and the airplane manufacturers 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 that may be 

applicable to the test. 

2. Brake Pressure Test 

Verify brake pressure tests conducted under the TSO are satisfactory 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 manufacturers 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 assure 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]. 
04/12/05 AC 23-17B 

81 

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 the landing performance is adequate to the previously approved 

performance data shown in the AFM. If the AFM performance data is not 

available because it is not required by the airplane certification basis, the 

manufacturers data (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 enough static torque when parked and during 

appropriate engine run up conditions [reference part 23, § 23.735(b)]. 

8. Brake Response Tests 

During the 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 relevant to endurance of landing gear system 

components. 

5. INSTRUCTIONS FOR CONTINUED AIRWORTHINESS 

A PMA applicant may be required to provide 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 acceptable for the applicants 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 people 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. 
AC 23-17B 04/12/05 

82 

For a part based on an STC, identifying installation eligible type certified products 

should include reference to the STC. Under 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 11: Because of 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. 

b. Part Numbering 

The PMA holders' part should be numbered so it is sufficiently different from the 

OEM holders' part number to be distinguishable. The OEM holders' part number with 

a prefix/suffix is enough 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 holders product 

line. The FAA-PMA letter should show the type approved part number with which 

the PMA holders' 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 with 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 

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. 
04/12/05 AC 23-17B 

83 

Amendment 23-7 and Subsequent

Amendment 7 was a major change in the standard. The following explanation in NPRM 67-

14 is stated as follows: “Part 23 does not provide for the determination of the energy 

absorption requirements of brakes. The present requirement in Sec. 23.735 (b) may not 

ensure the design of wheel brake combinations adequate for safe operation of the airplane. 

Proposed Sec. 23.735 (a) would therefore provide standards similar to the energy absorption 

determination requirements for transport category airplanes (Sec. 25.735 (f)). Proposed Sec. 

23.735(b) would reflect the fact that, for aircraft with high power-to-weight ratios, it may be 

impossible to prevent motion, with locked wheels, when takeoff power is applied to the 

critical engine. Some question has arisen as to whether such motion constitutes "rolling" of 

the airplane within the meaning of present (a). It is not the intent of present (a) to prevent 

this motion, but only to ensure that the brakes can prevent the wheels themselves from 

rolling. The proposed energy absorption requirements reflect current industry practice in 

the design of small airplanes.”

Amendment 23-42 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 (that is, brakes, autopilots, and ice protection systems). 

In the past, the Small Airplane Directorate has not considered brake failures catastrophic or 

severe-major because of mitigating factors, which include: 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)): 

In the following circumstances, the brake system and associated systems must be designed 

and constructed so 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), should be considered as connecting or transmitting elements unless it is 

shown 
AC 23-17B 04/12/05 

84 

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 airplanes.

Amendment 23-49 and Subsequent

The intent of the proposed revision to NPRM 94-21 is given as follows: “Section 23.735(a) 

would be revised to state plainly that wheel brakes must be provided. A proposed new 

Sec. 23.735(c) would require the brake system to be designed so that the brake 

manufacturer's specified brake pressures are not exceeded during the landing distance 

determination required by Sec. 23.75. Proposed new Sec. 23.735(e), applicable to commuter 

category airplanes, would require establishing the minimum rejected takeoff brake kinetic 

energy capacity rating of each main wheel brake assembly. Section 23.45 provides that the 

determination of the accelerate-stop distance for commuter category airplanes be made in 

accordance with the applicant's procedures for operation in service. The proposed 

requirement is needed to ensure that the brakes will perform safely under accelerate-stop 

conditions.” 

EASA AMC 23.735(c) is acceptable for FAA certification. 
04/12/05 AC 23-17B 

23.737 Skis

The corresponding rule in CAR 3 is CAR 3.364. 

There is no corresponding rule in the Airship Design Criteria.

Amendment 23-45 and Subsequent 

See § 23.505, Supplementary conditions for ski-planes, for additional guidance about 

aircraft skis. 

85 
AC 23-17B 04/12/05 

23.745 Nose/tail wheel steering

No policy available as of September 30, 2003. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

86 
04/12/05 AC 23-17B 

FLOATS AND HULLS 

23.751 Main float buoyancy 

The corresponding rule in CAR 3 is CAR 3.371. 

There is no corresponding rule in the Airship Design Criteria.

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 the aircraft remain afloat with any two compartments of the main floats 

flooded. The history does not support the position the aircraft will remain afloat 

indefinitely without capsizing with two compartments flooded. However, if an 

unsafe condition exists, certification should be denied under part 21, § 21.21(b)(2). 

An example of an unsafe condition could be capsizing so rapidly 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 AFM. 

Amendment 23-45 and Subsequent 

A revision to NPRM 90-18 explains this amendment as follows: “This proposal 

revises Section 23.751 to clarify the buoyancy requirements for main floats in 

paragraph (a)(1) by specifying an 80 percent excess in buoyancy for each main float 

above the buoyancy required by that float to support the maximum weight of the 

seaplane. Additionally, the words "without capsizing" are added to paragraph (a)(2)

to clarify the extent of flotation necessary after main float compartment flooding. 

A strict interpretation of existing Section 23.751(a)(1) results in a buoyancy excess of 

80 percent of the maximum weight of the seaplane when the design consists of only 

one main float, or a total of 180 percent of the maximum weight. However, on 

seaplanes having two main floats, each float would be required to have buoyancy of 

80 percent in excess of that necessary to support the seaplane, or 180 percent of the 

maximum weight of the seaplane; for a total of 360 percent of the maximum weight. 

For designs having three floats, each float would be required to support 180 percent 

of the maximum weight for a total of 440 percent. This is neither the intent of the rule 

nor the practice of industry. 

The change to paragraph (a)(2) is intended to clarify the fact that the seaplane be 

afloat in the upright condition.” 

87 
AC 23-17B 04/12/05 

23.753 Main float design

No policy available as of September 30, 2003. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

88 
04/12/05 AC 23-17B 

23.755 Hulls 

The corresponding rule in CAR 3 is CAR 3.372. 

There is no corresponding rule in the Airship Design Criteria.

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. 

89 
AC 23-17B 04/12/05 

23.757 Auxiliary floats

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.373. 

There is no corresponding rule in the Airship Design Criteria. 

90 
04/12/05 AC 23-17B 

PERSONNEL AND CARGO ACCOMMODATIONS 

23.771 Pilot compartment

No policy available as of September 30, 2003. 

The corresponding rule in CAR 3 is CAR 3.381. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.28. 

91 
AC 23-17B 04/12/05 

23.773 Pilot compartment view

The corresponding rule in CAR 3 is CAR 3.382. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section .29. 

Amendment 23-14 and Subsequent

See AC 23-8B, “Flight Test Guide for Certification of Part 23 Airplanes,” for 

guidance on this regulation. EASA AMC 23.773 references the CS-23 Flight Test 

Guide. 

Amendment 23-45 and Subsequent 

A proposed revision to NPRM 90-18 explains the rewrite of this section as follows: 

“This proposal is based in part on conference recommendation 272, on conference 

comments, and on a post conference review of the adequacy of previous 

certifications, which establishes a precedent for compliance with existing Section 

23.773. It is not the intent of this proposal to require windshield heat on all small 

airplanes, to preclude open cockpit designs or to prohibit the pilot from using a cloth 

to wipe the windows. It does, however, define requirements to assure that a means 

exists to remove or prevent the formation of fog or frost on the inside of the 

windshield, specifies the extent of credit to be given to pilot actions and defines the 

area of windshield and windows to be kept clear. 

Paragraph (a)(1) of this proposal requires an extensive, clear, and undistorted view 

sufficient to enable the pilot to perform any maneuvers within the operating 

limitations of the airplane, and specifies particular operations, such as taxi, takeoff, 

approach and landing to clarify the extent of view necessary for safe operation. 

Paragraph (b) of this proposal is included to address the condition where an airplane 

is operated at high altitudes, becomes cold-soaked, and is then descended into warm, 

moist air. Such conditions have resulted in the formation of frost on the inside 

surface of the windshield and crew compartment windows, which resulted in a limited 

or completely obscured view. Since, in such cases, compliance has been shown for 

the current Section 23.773(a)(3), a rule change is appropriate to address this 

condition. The FAA proposes to revise Section 23.773 to identify this condition and 

to clarify the extent of actions taken by the pilot to remove such moisture.” 

The Final Rule 26269 for Amendment 23-45 states as follows: “The FAA proposed 

to revise the 23.773 requirements for the pilot compartment view to address the 

environment expected in all the operations requested for certification. The JAA 

states that it will consider this change for JAR 23 but that it proposes to retain 

present paragraph (b) relating to night flight tests. The GAMA contends that the 

words 

92 
04/12/05 AC 23-17B 

"must be shown in all operations for which certification is requested," could be 

interpreted to mean that the same view must be provided for all operations. 

The FAA does not agree with the GAMA interpretation. Section 23.773(a) and (a)(1)

require the pilot compartment view to be sufficiently extensive, clear, and undistorted 

to allow the pilot to perform the various functions identified in this proposal. The 

word "sufficiently" is included because the FAA recognizes that the view needed for 

one operation may differ from the view needed for another. The intent is also shown 

by the words "sufficiently large" used in Section 23.773(b). The FAA adopts Section 

23.773 as proposed.”

93 
AC 23-17B 04/12/05 

23.775 Windshields and windows

The corresponding rule in CAR 3 is CAR 3.383. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.30.

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 for Federal Standard 406, 

Method 3022 or equivalent. 

Based on 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 criterion 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 MOCs 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. 

94 
04/12/05 AC 23-17B 

NOTE 12: We are not aware of any case where an applicant has used an analysis 

without testing, but we should consider this if an applicant proposes it. 

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. 

Amendment 23-7 and Subsequent 

A proposed revision to NPRM 67-14 explains this amendment as follows: “Sudden 

decompression of small cabins due to transparency failure is potentially catastrophic. 

Experience with transport category airplanes has shown that decompression has 

been averted in several cases because of the dual element design of transparencies. 

The value 25,000 feet is believed to represent a reasonable point of demarcation 

between the altitudes at which sudden decompression is and is not potentially 

catastrophic, considering the expected cabin sizes and operating altitudes of small 

airplanes.” 

The preamble material for Final Rule, Docket 8083 is explained as follows: “The 

notice proposed to add a new Sec. 23.775(e) imposing a fail-safe requirement for 

windshields, window panels, and canopies of airplanes certificated for operation 

above 25,000 feet. One comment objected, stating that the proposed requirement is 

unnecessary and an arbitrary dictation of design; that windshields should not be 

singled out over control systems or other components, and that transport category 

airplane decompression experience does not apply to Part 23 airplanes because the 

differential pressures are not the same. Commentator also suggested that if the 

proposal is adopted, the prescribed load should be the ultimate load. The FAA does 

not agree. Windshield strength cannot be controlled to the same degree of precision 

as is achieved with other materials, consequently redundancy is particularly 

important. Part 23 airplanes operate at the same altitudes as transport airplanes, 

where windshields have been lost and where windshield failure can be extremely 

hazardous. There is no justification for Part 23 design standards for windshields to 

be lower than those required in Part 25. The prescribed loading cannot be 

considered the ultimate load because a partially failed windshield would have 

essentially no factor of safety. Another comment objected to the proposal because 

paragraph (e) would require a fail-safe determination for all canopies of pressurized 

95 
AC 23-17B 04/12/05 

airplanes operated above 25,000 feet in addition to the testing already required 

under present paragraph (c). The intent of the proposal is to require airplanes 

certificated for operations above 25,000 feet to comply with the requirements of 

proposed paragraph (e) instead of the requirements of current paragraph (c), and 

paragraph (c) is amended to limit its applicability to pressurized airplanes that do 

not comply with the fail-safe requirements of paragraph (e). Except for this change 

the amendment is adopted as proposed.” 

Amendment 23-45 and Subsequent 

A proposed revision to NPRM 90-18 explains a change to this amendment as follows: 

“This proposal is intended to clarify the criteria for determining the cleared 

windshield area the FAA deems necessary to assure safe operation for icing 

certification. By specifically identifying the operational phases of takeoff, approach, 

landing, and taxi, this proposal is intended to prevent the past practice of certifying 

airplanes for operation in known icing conditions, with panels too small and too far 

in front of the pilot (in some cases, a single small panel centered on the windshield to 

be used from either pilot seat) to allow full operation of the airplane. In such cases, 

the runway is not always visible during approach when crosswinds result in large 

crab angles. Additionally, upon landing, the ability to locate and safely use taxiways 

is hampered because of the restricted view available to the pilot through the small 

panel. This proposal is not intended to preclude the use of such panels, but does 

identify the criteria for determining the size, location, and, if necessary, the number 

of panels. 

In addition, a proposal is made to require that the probable single failure of 

transparency heating systems not adversely affect the integrity of the airplane cabin. 

Such failures do occur and consideration of such occurrences is necessary as a 

minimum requirement for the type certification of new airplane designs.” 

EASA AMCs 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. 

96 
04/12/05 AC 23-17B 

23.777 Cockpit controls 

The corresponding rule in CAR 3 is CAR 3.384. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.31.

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. 

A proposed revision to NPRM 67-14 added a requirement under Amendment 23-7 

about fuel selectors. The explanation in the NPRM reads as follows: “Proper fuel 

management requires that the pilot be able to reach the fuel selector, at any time, 

without being forced to move the seat or the primary controls (which may be 

hazardous at low altitudes). Ability to reach the fuel selector under any condition 

requires that each selector be visible to the pilot with the seat in any normal inflight 

position.” 

Amendment 23-33 and Subsequent

This amendment added § 23.777(c)(4). It requires that airplanes with side-by-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 ELOS finding for airplanes with 

one set on the left hand side and one on or near the cockpit centerline. 

The preamble to Amendment 23-33 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 (SLPCs)

The purpose of this guidance is to provide recommendations for certification of 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 (Revolutions Per Minute (RPM) control), and mixture control 

(condition lever and fuel cutoff for turbine powered airplanes) into a single power 

97 
AC 23-17B 04/12/05 

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 (for example, §§ 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. 

Because of recommendations made by the NTSB, Amendment 23-33 provides 

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 as follows: “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, the basis for Amendment 23-33, describes 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, obviously 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 improve 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 ELOS 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. Sometimes, however, the applicable airworthiness standards may not 

98 
04/12/05 AC 23-17B 

be adequate because of 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. 

We have introduced 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 AC may be used for certification of an SLPC 

on part 23/CAR 3 airplanes. 

99 
AC 23-17B 04/12/05 

23.779 Motion and effect of cockpit controls 

The corresponding rule in CAR 3 is CAR 3.384. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.32.

Original Issue—Reserved 

The original rule only specified primary aerodynamic and throttle motion. 

Amendment 23-33 and Subsequent

This amendment requires the propeller control should move forward to increase 

RPM. Therefore, a propeller pitch control on the vertical instrument panel does not 

comply. A switch found in a horizontal position with forward motion to increase 

RPM does comply. Other designs would have to be considered by an ELOS finding. 

We would consider a propeller pitch control switch that would increase propeller 

RPM when moved to the “UP” position with a placard to show 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. 

100 
04/12/05 AC 23-17B 

23.781 Cockpit control knob shape

The corresponding rule in CAR 3 is CAR 3.384. 

There is no corresponding rule in the Airship Design Criteria.

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. 

101 
AC 23-17B 04/12/05 

23.783 Doors 

The corresponding rule in CAR 3 is CAR 3.389. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.33.

Original Issue and Subsequent

An ELOS finding process for part 23, § 23.783(b), should include: 

a. Pilot operated locks when the propeller stops turning. 

b. A special operating procedure to assure 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 inspect each individual lock of the locking mechanism. Means 

that do not allow 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. 

102 
04/12/05 AC 23-17B 

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. 

Amendment 23-36 and Subsequent

Section 23.783(c) was adopted to provide standards that would assure the opening 

means of passenger and crew doors were simple, easy to find, and could be operated 

in darkness. It also assured 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 the opening means of 

passenger and crew doors were kept as simple as possible, and that these doors could 

be found and opened in a timely manner. Paragraph 23.783(c)(4) was adopted 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 show 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) 

103 
AC 23-17B 04/12/05 

Pressurized cabin loads (for flight, § 23.365(a) and (b); and for landing, § 23.365(a) 

and (c)) 

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 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. 

EASA AMC 23.783(b) is acceptable for FAA certification. 

104 
04/12/05 AC 23-17B 

23.785 Seats, berths, litters, safety belts, and shoulder harnesses 

The corresponding rule in CAR 3 is CAR 3.390. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.34.

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 

wither by an 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 STC or field approval. 

However, the FAA should not forbid the approval of a retrofit shoulder harness 

installation as a minor change in: 1) the front seats of those small airplanes 

manufactured before July 19, 1978, and 2) 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 either 14 CFR, part 23 before Amendment 

23-20, part 3 of the CAR, 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: 

AC 43.13-2A, “Acceptable Methods, Techniques, and Practices Aircraft Alterations.” 

AC 21-34, “Shoulder Harness-Safety Belt Installations.” 

AC 23-4, “Static Strength Substantiation of Attachment Points for Occupant Restraint 

System Installations.” 

105 
AC 23-17B 04/12/05 

Installations approved as a minor change may not provide the occupant with the 

protection required either by 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 

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 Inspectors 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 TC 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 the occupant be protected from the 

ultimate inertia forces specified in § 23.561(b)(2)]. Section 23.785(g) also provides 

the following: “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.” 

Part 91 operating rule stems from § 23.2, Special retroactive requirements, 

Amendment 23-32, effective December 12, 1985. 

106 
04/12/05 AC 23-17B 

3. Belts or harnesses provided for in the design. 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 and 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 the airworthiness 

regulations have been met. The STC approvals are issued by the FAA ACOs. STC 

approvals are usually obtained by a shoulder harness installation kit supplier for 

multiple airplane installations in an airplane model or model series. 

ACs 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 ACs are also applicable to installations in 

airplanes having a certification basis of predecessor regulations (for example, 

CAR 3). 

107 
AC 23-17B 04/12/05 

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 manufacturers 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 consider 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. 

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. 

108 
04/12/05 AC 23-17B 

In 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 because of the tube geometry, the load at the aft harness attachment 

caused a tension in the rear spar carry through tube. The front seat shoulder harness 

upper end was attached to the rear spar carry through 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 provide instructions for 

continued airworthiness, prepared in accordance with § 23.1529. 

An STC cannot 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: 

- 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 STCs. 

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 (ASI). Field approvals are appropriate for alterations that involve little or 

no engineering. If the installation requires structural modifications, an ACO will 

need to help 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 show 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. 

109 
AC 23-17B 04/12/05 

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: 

- ACs 43.13-1A and 43.13-2A, as amended* 

- Manufacturers technical information (for example, manuals, bulletins, and kits) 

- FAA Field Approvals” 

*NOTE 13: AC 43.13-1B, Change 1, dated September 27, 2001 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 either have a 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 carry 

through tube, or other unsupported member, should not receive a field approval 

without a design approval from an ACO. 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 carry through member installation. This is in the 

center of the carry through member that is a hat section as found in metal skinned 

airplanes. Part of the figure shows the hat section is riveted to sheet metal skin 

(which would provide longitudinal support 

Personnel performing the field approval must assure 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 

as follows: “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.” 

110 
04/12/05 AC 23-17B 

Information provided to us by the Anchorage ACO suggests 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 carry through 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 the harness installation had saved their lives. 

In general, shoulder harness installations should not use the center of an 

unsupported wing carry through 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 carry through tube. 

It is acceptable for the carry through structure to be damaged in an emergency 

landing. However, it is unacceptable for the tube to fail in-flight. Carry through 

tubes, loaded in compression, may experience a beam-column buckling failure if 

the occupant applies a load to the shoulder harness attachment. Occasionally, small 

loads on the shoulder harness attachment may cause beam-column 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 SAE 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: 

- harnesses that meet TSO-C114 or Military Specification (MIL-SPEC) 

requirements, 

- harnesses that have been produced under a 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

111 
AC 23-17B 04/12/05 

the structural test and any associated structural analysis. Also, there is a possibility 

the airframe may be damaged during the pull test. In such installations, a pull test 

would not be done and there is no assurance the installation will provide occupant 

protection to the ultimate inertia force requirements (the 9.0-g forward force) of 

§ 23.561 or CAR 3.386. 

About 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. 

CAR 3.652, Functional and installation requirements, states as follows: “Each item

of equipment which is essential to the safe operation of the 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 either 

for 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 Belt Installations,” AC 21-25A, 

“Approval of Modified Seats and Berths,” and AC 43.13-2A, “Acceptable Methods, 

Techniques, and Practices—Aircraft Alterations.” 

112 
04/12/05 AC 23-17B 

Aft-facing and side-facing seats 

For aft-facing seats, seat obliqueness should be limited to 15 degrees unless 

additional occupant protection for side-facing seats is installed. 

Part 23 allows 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 as well 

as 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 degrees 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 both 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 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. The sidefacing 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, TC 

approval or STC approval of sport parachute modifications to small airplanes. The 

Small Airplane Directorate has received NTSB recommendations on 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. 

ACOs should not provide engineering assistance for sport parachute 

113 
AC 23-17B 04/12/05 

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. 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 

the crashworthiness of the airplane can be unacceptable or that an unsafe condition 

can be created by a modification. 

The small airplane certification basis is mostly CAR 3 or early part 23. The 

modifications include: 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 CG changes. 

Seating capacity increases can be acceptable if: 1) all people are protected in 

emergency landing loads per § 23.561 (3.386), 2) all persons are secured against 

sliding backwards in climbs to maintain an acceptable CG position, and 3) 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 people 

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 

114 
04/12/05 AC 23-17B 

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. ACOs 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. So far, 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 by an 

approved procedure to verify its relevance in the emergency landing loads. 

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 for 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 14: 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 AFM 

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 

(FMS) 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. 

115 
AC 23-17B 04/12/05 

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 

A revision to NPRM 67-14 explains this amendment as follows: “Experience 

indicates that a significant reduction in injuries and fatalities in small airplane 

accidents may be obtained by the installation and use of effective upper body 

restraints or by designing airplane interiors to either eliminate injurious objects 

within striking radius of the head or provide energy absorbing support for the upper 

torso. This proposal would therefore apply, to small airplanes, standards identical 

with those long administered for large airplanes under Sec. 25.785(c).” 

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. 

Amendment 23-36 and Subsequent

A revision to NPRM 86-19 explains this amendment as follows: “The FAA is 

proposing a substantial revision of the current and new requirements for seats, 

berths, litters, safety belts, and shoulder harnesses. The revision is considered 

necessary to present the proposed requirements in a more logical sequence as a 

result of the new requirements being proposed by this rulemaking action. The new 

requirements are based to a large extent on the proposals submitted by the GASP to 

the Part 23 Airworthiness Review Program. The GASP proposals were based in 

large measure on FAA, NASA, and NTSB research studies and impact/accident 

analyses.” 

See guidance for § 23.562, Emergency landing dynamic conditions. 

See guidance for dynamic seat certification in AC 20-146, “Methodology for 

Dynamic Seat Certification by Analysis for Use in Part 23, 25, 27, and 29 Airplanes 

and Rotorcraft.” 

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 

116 
04/12/05 AC 23-17B 

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. This is 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. 

117 
AC 23-17B 04/12/05 

23.787 Baggage and cargo compartments 

The corresponding rule in CAR 3 is CAR 3.392. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.35.

Original Issue and Subsequent

Questions have been raised about § 23.787(c), which requires an ultimate inertia 

forward force of 4.5g for to protect passengers from any cargo compartment. This 

regulation is related to the CAR, which imagined 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 satisfactory 

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 assure 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 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 

118 
04/12/05 AC 23-17B 

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 ATCs or 

STCs specific to this operation. 

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. This is 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 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 show 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 FSDO. 

Amendment 23-14 and Subsequent 

See AC 23-2, “Flammability Tests.” 

A proposed revision to NPRM 71-13 explains this amendment as follows: “This 

proposal would make it clear that it is not only the cargo shifting that would affect 

the CG of the airplane that is covered by the rule but it is also cargo shifting which 

could damage essential controls, equipment etc. In addition, for airplanes in which 

cargo is carried in the passenger compartment the proposal would make the higher

119 
AC 23-17B 04/12/05 

inertia forces of Sec. 23.561(b)(2) applicable to the protection of the passengers.” 

Amendment 23-36 and Subsequent

A proposed revision to NPRM 94-21 explains the amendment as follows: “Section 

23.787(c) is proposed to be revised because, as presently stated, the required 

ultimate forward inertia force for cargo restraint is not adequate considering the 

current requirements of Section 23.561. It was the consensus at the conference that 

cargo restraint should be at least to the ultimate inertia forces of Section 23.561 to 

adequately protect occupants forward of the cargo. In addition, when designs 

provide for cargo to be carried in the same compartment with occupants, it is 

proposed that means be provided to restrain the cargo, at least to the loads resulting 

from the emergency landing dynamic conditions being proposed in Section 

23.562(b)(2). It is considered necessary to protect occupants from cargo being 

forced into their occupied area as a result of an emergency landing when they are 

otherwise being adequately protected from serious injury. The increased ultimate 

static load factors will achieve this objective.” 

The rigid movable/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). 

Before 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 

movable/removable cargo restraint bulkhead in the cabin. The reason was to bring 

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. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this amendment as follows: “Section 

23.787 would be revised by extending the present requirements for cargo 

compartments to baggage compartments. As proposed, future baggage 

compartments on all airplane categories would be required to be placarded for their 

maximum weight capacity; have a means to prevent the baggage from shifting; and 

have a means to protect controls, wiring, lines, and equipment or accessories that are 

located in the compartment and whose damage or failure would affect safe operation 

of the airplane. These standards have been applicable to cargo compartment design

120 
04/12/05 AC 23-17B 

for same time and should be applied to baggage compartments since the same safety 

factors are involved. Because manufacturers recognize the need for these standards, 

many of these provisions have been included in the current design of baggage 

compartments and, therefore, the proposed requirements are not expected to create a 

significant burden. With this revision the commuter category requirements of Sec. 

23.787(g) would be redundant and that requirement is being removed. 

Proposed revisions to this section would also move the substance of paragraphs (d)

and (f) to a proposed new Sec. 23.855, which will address cargo and baggage 

compartment fire protection. 

Proposed new paragraph (c) of this section would require flight crew emergency 

exits on all cargo-configured airplanes to meet the requirements of Sec. 23.807. This 

requirement would provide increased assurance that flight crews of all cargo 

airplanes will have ready access to an emergency exit.” 

121 
AC 23-17B 04/12/05 

23.791 Passenger information signs

No policy available as of September 30, 2003. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

122 
04/12/05 AC 23-17B 

23.803 Emergency evacuation 

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

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. 

It is 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. This is when 

certification of emergency exits is done per § 23.807(d)(4). AC 20-118A is still 

applicable except for paragraph 5a(3)(vi). 

123 
AC 23-17B 04/12/05 

23.805 Flight crew emergency exits

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.36.

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. 

Section23.807(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 rigid 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.” 

124 
04/12/05 AC 23-17B 

23.807 Emergency exits 

The corresponding rule in CAR 3 is CAR 3.387. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.36.

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. Thus, 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 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 shown. 

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. On the acrobatic category, if the canopy is not 

jettisonable, it should be shown that the canopy could 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 would 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. 

On doors between the pilots' compartment and the passenger compartment that are 

likely to block the pilots' egress in a minor crash landing, there should be an exit in 

the pilots' 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 

section 23.807(a) below or by the conditions for acceptance of rigid doors in section 

23.807(b) below. Rigid doors are only acceptable by placarding the doors to be 

125 
AC 23-17B 04/12/05 

latched in the open position for takeoff and landing, providing the conditions in 

section 23.807(b) below are in compliance. 

a. The purpose of the test is to show the door between the pilots' compartment and 

the passenger compartment will not block the pilots' escape if the door is jammed. 

Acceptable means of compliance is by showing the door is frangible and the flight 

crew can egress the airplane without help 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 people with no particular escape abilities. The estimated 

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 (fifth to 

95th percentile). The female participant will break the door and be the first 

person through the exit without help from the male participant. Instructions 

for improving the egress should be limited to those instructions that are 

provided in either the FAA approved AFM or on related placards, or both. 

(3) Determine 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 told 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 the door is frangible should be placarded 

on the door(s) and should be noted in the limitations section of the AFM.

126 
04/12/05 AC 23-17B 

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 on 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 sought, the pilots' compartment—with 

the doors open—should be free from glare and reflections that could interfere 

with a pilots' vision, per § 23.733. 

(5) The doors should be placarded in accordance with § 23.1557. The placards 

should state the doors are to be latched in an open position before takeoff and 

landing. Placard information should be in the imitations section of the AFM. 

The following should be considered when approving STCs for cargo conversions (no 

passenger seating) on airplanes with a certification basis of 14 CFR, part 23, 

including commuter category airplanes: 

• Unless otherwise mentioned, 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 pilots 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). 

127 
AC 23-17B 04/12/05 

• It is not acceptable to remove 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 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, when consideration is given 

to the effect of a post-crash environment. An emergency evacuation 

demonstration, therefore, is not acceptable as an ELOS since it is not 

meant to allow alleviation or deviation of specific requirements. 

If the airplane is configured for passenger and cargo (Combi), the requirements of 

this section should be met for the passenger compartment. In addition, cargo should 

be located: 1) so it does not obstruct either access to or use of any required 

emergency or regular exit; 2) so it does not obstruct the use of the aisle between the 

crew and passenger compartment; and 3) so it meets the additional requirements of 

§ 135.87. 

Amendment 23-7 and Subsequent 

A proposed revision to NPRM 67-14 explains this amendment as follows: “Section 

23.807 (a) (1) exempts all airplanes with five-or-less occupants from the requirement 

to have an emergency exit opposite the main door. For airplanes with engines 

mounted on the wings or on the side of the fuselage, the main door may be blocked by 

engine fire regardless of the number of occupants. Therefore the five-or-less 

exemption would be discontinued for these aircraft but would be continued for single 

engine airplanes and centerline thrust twins.” 

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

A proposed revision to NPRM 83-17 added Amendment 23-34. “The source for this 

revision is part 135, Appendix A., sections 32(a) and (b); SFAR 41, section 5(e), 

Doors and Exits, paragraphs (g), (i), and (j)”. 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) on 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 

128 
04/12/05 AC 23-17B 

operation of the exit even in darkness, and § 23.811(b) requires illuminating the exit 

sign. As an added safety feature, it is recommended the operating handle be selfilluminated and marked with a red arrow and the word “OPEN” in red letters placed 

near the head of the arrow. If necessary, other relevant 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 for 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. 

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: 1) under the inertia forces 

of § 23.561 and 2) following the collapse of one or more legs of the landing gear. A 

demonstration of this failure mode is beyond the intent of this rule. It should be 

shown by orthographic drafting techniques or test (that is, 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 ELOS 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 as well as 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 enough 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: 1) 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 2) the most critical dimension, width or height, greater 

than 19 inches, but lacking the shape to admit a 19 x 26 inch ellipse. 

129 
AC 23-17B 04/12/05 

Acceptable Means of Compliance 

Alternatives for compliance to the airworthiness standards are allowed by an ELOS. 

One method for determining compliance by an ELOS is by the test procedure below. 

Demonstrations have shown the emergency exit size and shape affect the time and 

ease of an emergency evacuation. An ELOS should only be considered if the exit 

meets the logical limits that 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. 

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 removed 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 preserve 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 2) 

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 either 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 allowed to help participants in 

descending from a wing when over-wing exits are used if the acceptance rate of 

130 
04/12/05 AC 23-17B 

the ramp or stand is no greater than that of the assist means of the airplane in a 

crash landing situation. Mats may be used on the floor or ground to protect 

participants. No other equipment that is not part of the airplanes 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 allowed any “practice” runs, but they may be briefed on 

the purpose of the test to show a rapid emergency evacuation of the airplane. 

They should not be briefed that the test is to compare exits. An example of an 

acceptable 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 for 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. This is 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 for physical agility, age, sex, 

and weight. This can be achieved by first dividing the group by age and 

131 
AC 23-17B 04/12/05 

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 removes the effects of an individuals learning, fatigue, and agility. 

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 time can 

be recorded or synchronized electric cameras may be used with the time 

superimposed in the film processing. A signal light to show the beginning and 

end of each trial should also be arranged in 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 may be possible one group contains 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 events 

to be evaluated and will aid 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.

132 
04/12/05 AC 23-17B 

Type Certificate Data Sheet (TCDS). An equivalent level of safety should be part 

of the type certification basis and noted on the TCDS. 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.” 

Amendment 23-36 and Subsequent 

This amendment is explained by Notice 86-19 as follows: “The issue of number and 

location was addressed by one proposal to require that for all airplanes with a 

seating capacity of two or more, excluding airplanes with canopies, have at least one 

emergency exit on the opposite side of the cabin from the main door specified in 

Section 23.783. There have been egress difficulties experienced with center line 

engine airplanes with a seating capacity of two or more, excluding those airplanes 

with canopies, and no emergency exit or door opposite the main cabin door. In some 

cases, occupants have had to kick out windows to egress the airplanes after 

survivable accidents.” 

Amendment 23-46 and Subsequent 

A proposed revision to NPRM 90-20 explains this amendment as follows: “For 

standardization between the airworthiness requirements for commuter category and 

transport category airplanes, and to simplify the alternate emergency exit standards 

proposed in this notice, the FAA proposes to move certain requirements of Sec. 

23.807(d)(1) to a new paragraph (f) in Sec. 23.783 and proposes a new paragraph 

(d)(3) in Sec. 23.807. The standards in proposed Sec. 23.783(f) require that each 

passenger entry door qualify as a floor level emergency exit and provide 

requirements for integral stairs when installed at a passenger entry door. The 

standards in proposed 23.807(d)(3) require that each emergency exit that is not a 

floor level exit either be located over the wing or, if not less than six feet from the 

ground, have an acceptable means to assist the occupants in descending to the 

ground. There are no substantive differences between the requirements in proposed 

Sec. 23.783(f) or in proposed Sec. 23.807(d)(3) and those requirements proposed to 

be removed from Sec. 23.807(d)(1).” 

A proposed revision to the NPRM explains the following: “This notice proposes new 

requirements for emergency exit ditching provisions for multiengine airplanes that 

are type certificated to the airworthiness standards of part 23. The FAA anticipates 

an increase in the use of commuter category airplanes and multiengine normal 

category airplanes in over water operation. Airports developed near large bodies of 

water increase the number of departures and approaches that are conducted over 

water. Since ditching provisions may be critical for occupant egress following an 

emergency landing in water, the standards in proposed Sec. 23.807(e) ensure the 

availability of exits for emergency egress following an emergency landing in water.”

133 
AC 23-17B 04/12/05 

Amendment 23-49 and Subsequent

The following flow chart is valid for emergency exit in normal, utility, and acrobatic 

category airplanes. 

Non-Commuter Emergency Exit Flowchart

One Emergency Exit on Opposite Side

of Cabin from Main Door

Passenger Exits

Available to Crew

One Emer. Exit on Each Side of A/C

OR Top Hatch

IN CREW AREA

Pilot’s Compartment Separated

from Passenger Compartment 

by a door that will likely

block Pilot’s Escape

Require One Exit in Pilot’s Compartment

Exits near a Hazard

RECONFIGURE

Multiengine A/C

One Exit Above WL on

Each Side of Aircraft

OR

Overhead Hatch

STOP

START

A

A

No

Yes

Yes

No

Yes

No

No

Yes

23.807(a)(1)

23.805(a)

23.805(a)

23.807(a)(3)

23.807(a)(3)

23.807(a)(4)

23.807(a)(4)

23.807(e)(1)

23.807(e)(1)(2)

134 
04/12/05 AC 23-17B 

23.811 Emergency exit marking 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.36.

Amendment 23-36 and Subsequent

For small airplanes with emergency exits that open 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 prominent manner. 

c. Outline the exit handle with a band of a contrasting color. 

d. Mark the exit with any other prominent 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, 

except that they must be obviously identifiable from outside the airplane. Section 

23.1557(d) specifically requires that each placard and operating control for each exit 

must be red. Also, a placard must be near each emergency exit control and must 

clearly show the location of that exit and its method of operation. 

Since the emergency exit exterior markings are obviously identifiable from the 

outside of the airplane, to further require red placards that show the location of these 

exits does not make sense. There might be instances where the red placard may not 

be clearly identifiable because of the exterior color of the fuselage. Section 23.807 

gives further guidance on 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.

135 
AC 23-17B 04/12/05 

Amendment 23-46 and Subsequent 

The explanation for this amendment in NPRM 90-20 states as follows “The 

emergency exit marking standards for transport category airplanes, as stated in Sec. 

25.811, have specific requirements that go beyond the current commuter category 

airplane emergency exit marking standards. Therefore, this notice proposes to add a 

new Sec. 23.811(c) providing additional airworthiness requirements for emergency 

exit marking that would be applicable when certification to the emergency exit 

provisions of Sec. 23.807(d)(4) is requested. The standards in proposed Sec. 

23.811(c)(1), which are similar to the requirements of Sec. 25.811(a), ensure 

conspicuous marking for each emergency exit, its means of access and its means of 

opening for rapid identification and operation of the exits in an emergency condition. 

The standards in proposed Sec. 23.811(c)(2), which are similar to the requirements of 

Sec. 25.811(b), ensure that the airplane occupants can readily identify and locate the 

emergency exits on the opposite side of the cabin from where they are seated. The 

standards in proposed Sec. 23.811(c)(3), which are similar to the requirements of 

Sec. 25.811(c) that ensure that the airplane occupants can locate the emergency exits 

when the cabin is filled with dense smoke. The standards in proposed 

Sec.23.811(c)(4), which are similar to the requirements of Sec. 25.811(e)(1), ensure 

that the operating handle and the instructions for opening the emergency exits are 

shown by a marking that is readable from a distance of 30 inches. The standards in 

proposed Sec. 23.811(c)(5), which are similar to the requirements of Sec. 

25.811(e)(2), ensure that there is sufficient lighting to allow identification of the 

passenger entry door operating handle. The standards in proposed Sec. 23.811(c)(6),

which are similar to the requirements of Sec. 25.811(e)(4), ensure the ease of access 

and operation of a passenger entry door with a locking mechanism that is released by 

a rotary motion of the handle. The standards in proposed Sec. 23.811(c)(7), which 

are similar to paragraphs (1) and (2) or Sec. 25.811(f), ensure that the emergency 

exits are externally marked so that they can be readily identified in conditions of low 

lighting or poor visibility. These proposed requirements would result in emergency 

exits that are easier to locate and open in adverse conditions.” 

136 
04/12/05 AC 23-17B 

23.812 Emergency lighting

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-46 and Subsequent 

The explanation for this rule adopted in Amendment 23-46 states as follows: “This 

proposal adds requirements for an emergency lighting system, to be applicable when 

an applicant for type certification chooses to comply with the alternate emergency 

exit provisions of proposed Sec. 23.807(d)(4) of this part. The proposal defines 

specific minimum requirements for supplying power, arming, and activating the 

emergency lighting system. The impact-activation requirement is consistent with that 

for emergency locator transmitters. The proposal also includes illumination, 

function, and survivability requirements for the emergency lighting system. An 

emergency lighting system complying with these proposed requirements would aid 

occupants in locating the emergency exits and getting to those exits after an 

emergency landing.” 

137 
AC 23-17B 04/12/05 

23.813 Emergency exit access 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-36 and Subsequent 

A change to this amendment was added by Final Rule docket 25147 as follows: “In 

addition, the requirements of Section 23.807(d)(4) are moved to a new Section 

23.813, titled Emergency exit access", in this final rule.” 

See § 23.807 Means of Compliance for policy on acceptance of seatbacks obstructing 

emergency exits that can be easily moved. 

Amendment 23-46 and Subsequent 

The explanation in NPRM 90-20 for this amendment states as follows: “This 

proposal adds requirements to ensure emergency exit accessibility, to be applicable 

when an applicant for type certification choose to comply with the alternate 

emergency exit provisions of proposed Sec. 23.807(d)(4) of this part. Structural 

failures or yielding of the airframe can occur during an emergency landing or minor 

crash event that may result in one or more emergency exits or the passenger door 

becoming unopenable. Since the total number of exits available for emergency 

egress can be fewer with the alternate emergency exit requirements, this proposal 

defines minimum unobstructed aisle width at the passenger entry door and adds other 

requirements to ensure that any partitions or doorways within the passenger 

compartment will not hinder occupant access to the exits during an emergency 

situation.” 

138 
04/12/05 AC 23-17B 

23.815 Width of aisle 

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Amendment 23-34 and Subsequent

This rule, adopted “from SFAR 41, Sec. 5(e), Doors and Exits, paragraph (k), 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. 

Amendment 23-46 and Subsequent 

A proposed revision to NPRM 90-20 explains this amendment as follows: “This 

proposal requires increased aisle widths to be applicable when an applicant for type 

certification chooses to comply with the alternate emergency exit provisions of 

proposed Sec. 23.807(d)(4) of this part. The proposed increased aisle width 

requirements are intended to ensure that the airplane passengers can reach an exit in 

an emergency situation even though the floor structure has been warped or there are 

seats or other items protruding into the normal aisle space.” 

139 
AC 23-17B 04/12/05 

23.831 Ventilation 

The corresponding rule in CAR 3 is CAR 3.393. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.38.

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 either 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 

five 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 on allowable carbon monoxide (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. This is done 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 

AC 20-32B. These are acceptable either for a TC applicant, an STC 

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

The source for this revision is shown by NPRM 83-17 as “Part 135, Appendix A, 

section 60, and SFAR No. 41, section 4(c)(3), which incorporates § 25.831(d) by 

reference”. For pressurized commuter category 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 AFM, Emergency or Non-Normal 

(Abnormal) Procedures Section, or on approved placards. 

140 
04/12/05 AC 23-17B 

141 
AC 23-17B 04/12/05 

1. What is the acceptable level of smoke, in the cockpit, as a result from this type of 

test? 

The test to show compliance to § 23.831(b) is for "all" fires that are possible not just 

engine fires. Conservatism requires the test begin with smoke sufficient to obscure 

the instrument panel. The plane must remove this smoke without depressurizing the 

plane below safe limits per § 23.831(b). 

2. What is the acceptable level of toxicity of the smoke? 

FAA has a requirement for carbon monoxide in § 23.831(a), but no other material. 

The applicant must determine what combustion by-products will be generated from

the reasonably probable failure of an engine fire in a reciprocating engine. 

3. Three minutes or five minutes for clearing smoke from the cockpit? 

The three-minute period is acceptable not to remove all smoke but to permit the pilot 

to safely fly the plane (see the flight instruments). 

4. What does "readily accomplished" mean? 

This rule came from part 135, Appendix A, which was based on § 25.831(d). It 

means the flight crew can easily activate the depressurization system while seated and 

belted in the cockpit. In a single pilot airplane, this means from the left seat. 

5 What does "without depressurizing beyond safe limits," mean?

The source for this rule was the same as question 4. Part 91 specifies that a flight 

crew can operate at 14,000 feet cabin pressure for 30 minutes. This is the maximum

level the plane can be depressurized without supplemental oxygen for operation 

above 14,000 feet for more than 30 minutes. However, applicants can use Time of 

Useful Consciousness for an emergency descent as in a pressurization failure. They 

must allow time for the cockpit to clear of smoke sufficiently for the flight crew to 

safely begin the descent to 14,000 feet or below. 

6. Can the smoke evacuation test be done on the ground?

Yes, a ground test is acceptable IF the applicant can clearly show that all relevant test 

boundary conditions match the respective flight conditions (such as airflows, and 

pressure fields.). If an applicant elects to have a cabin altitude limiter, the ground test 

must be conducted at or above the setting of the limiter (example: test site above 

14,000 feet Mean Seal Level (MSL), if the limiter is set at 14,000), otherwise no. 

142 
04/12/05 AC 23-17B 

7. Can the applicant assume an emergency descent? 

For answer 1, the smoke must obscure the panel. An emergency descent cannot be 

started until the smoke evacuation takes effect.

143 
AC 23-17B 04/12/05 

Amendment 23-42 and Subsequent 

This amendment makes paragraph (b) and the policy under Amendments 23-34 and 

subsequent applicable to all part 23 pressurized airplanes, not just commuter 

category. 

AC 25-9A, “Smoke Detection, Penetration, and Evacuation Tests and Related Flight 

Manual Emergency Procedures,” dated January 6, 1994, provides an acceptable 

means of compliance for the smoke evacuation requirements of this section. 

144 
04/12/05 AC 23-17B 

PRESSURIZATION 

23.841 Pressurized cabins 

The corresponding rules in CAR 3 are CAR 3.394 and 3.395. 

There is no corresponding rule in the Airship Design Criteria.

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 

the cabin doors and emergency exits can be opened and evacuation is not impaired. 

Time to equalize the ambient and cabin pressures should be shown. Inflatable door 

seals, if installed, are subject to the requirements of this rule. 

Paragraph (f) of § 23.841 requires a warning device for safe or preset pressure 

differential and absolute cabin pressure. A warning is interpreted to suggest 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. 

Amendment 23-7 and Subsequent 

This amendment changed the warning requirements in paragraph (f). The 

explanation in NPRM 67-14 states as follows: “Present Sec. 23.841 (f) does not 

describe what pressure warning devices are acceptable. This proposal would follow 

the approach of Sec. 25.841 (b) (6), that is, that instrument markings are sufficient 

for safe warning of pressure differential limits, whereas aural or visual means are 

necessary for safe warning of absolute pressure limits.” The Final Rule, Docket 

8083, added the following: “The proposed requirement permits the use of markings 

on the presently required pressure differential indicator as the warning indicator.”

145 
AC 23-17B 04/12/05 

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 proper 

to use an emergency descent procedure to show compliance to this rule when 

compliance can be achieved through design. The Emergency Operations Section of 

the AFM should include an emergency descent procedure for use in a rapid 

decompression from any failure not withstanding the likelihood of its occurrence.

Amendment 23-17 and Subsequent

A proposed revision to NPRM 75-10 explains Amendment 23-17 as follows: “This 

proposal would make clear that the rule applies to cabin pressure and the rate of 

change of cabin pressure. The proposal would also conform the rule to current 

practice, which is to provide a warning when the cabin pressure altitude exceeds 

10,000 feet. 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 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 the CAR for 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.

146 
04/12/05 AC 23-17B 

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. Because of 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. This was because most of the cabin pressure control 

system design for small airplanes drew heavily on 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). 

(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) for 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) Because of 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 safety-dump 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 

because of cabin pressure differential can be reduced accordingly. 

(iii) When a threatening cabin overpressure condition exists because of 

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 and fumes,, the 

cabin safety-dump valve may be used, depending on the conditions,.

147 
AC 23-17B 04/12/05 

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 relevant part, 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 the warning for cabin 

pressure would meet the applicable requirements if it occurred when 

cabin absolute 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 MSL is 

exceeded. The preamble for this change indicated this proposal was 

adopted because most 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 because of.

148 
04/12/05 AC 23-17B 

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 

assigning 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. 

Air Conditioning System R12 to R134 conversion

Direct conversion without certification is not possible, because of the 

characteristics of the refrigerants. Incompatibilities can include o-rings, gaskets, 

washers, receiver-dryers, expansion valves, and hoses. The system oil must also 

be changed, so a flush is required. 

Amendment 23-49 and Subsequent 

This amendment changed the 33,000 feet in paragraph (a) to 25,000 feet based on 

European Joint Aviation Requirements Proposals. 

149 
AC 23-17B 04/12/05 

23.843 Pressurization tests

Original Issue and Subsequent

The corresponding rule in CAR 3 is CAR 3.396. 

There is no corresponding rule in the Airship Design Criteria. 

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. 

150 
04/12/05 AC 23-17B 

FIRE PROTECTION 

23.851 Fire extinguishers

The corresponding rule in CAR 3 is CAR 3.388. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.39.

Amendment 23-34 and Subsequent 

A proposed revision to NPRM 83-17 lists the source for Amendment 23-34 as SFAR 

41, sections 7(d) and (e). 

See AC 20-42C, “Hand Fire Extinguishers for Use in Aircraft,” for guidance. 

European Aviation Safety Agency (EASA) AMC 23.851(c) accepts the AC as an 

AMC pending research into Halon replacement. 

Amendment 23-45 and Subsequent 

A proposed revision to NPRM 90-18 explains this amendment as follows: “This 

proposal extends the commuter category requirement for a hand fire extinguisher in 

the pilot compartment to all small airplane categories. Additionally, this proposal 

provides minimum acceptable standards for on-board hand fire extinguishers.”

“Besides this specific rule for fire extinguishers, there are other rules including 

§ 23.561, Emergency Landing Conditions (extinguishers cannot be a missile in an 

emergency landing); § 23.303, Factor of safety; § 23.601, Design and Construction, 

General; § 23.603, Materials and workmanship; and § 23.613, Material strength 

properties and design values. You will need to contact an Aircraft Certification 

Office for an installation in a small airplane.” 

151 
AC 23-17B 04/12/05 

23.853 Passenger and crew compartment interiors 

The corresponding rule in CAR 3 is CAR 3.388. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.39.

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 as 

follows: 

a. The only materials for which flammability testing is not required are those that 

would not contribute significantly to propagating a fire per 14 CFR, part 23, § 23.853. 

These materials/parts are typically small parts. Material certification by an applicant 

or their supplier can be used in the determination the material will not significantly 

contribute to propagating the fire. Company/supplier material certifications cannot 

be used instead 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. 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 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 DER who has authorization to witness a test. We will also accept from

a DER submittal of an FAA Form 8110-3 that certifies 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 15: 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 

152 
04/12/05 AC 23-17B 

quality assurance report. When the material supports a certification project or an 

alteration or

153 
AC 23-17B 04/12/05 

repair, FAA Form 8110-3 is the DERs 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 acceptable, the DER may generate an FAA Form 8110-3 that references the 

test report. For 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 Normal Flame resistant if smoking allowed 

(1945-1946) Restricted 

Effective Experimental 

November 13, 1945 

CAR 3 Normal, Flash resistant 

(1946-1949) Utility and 

Effective Acrobatic 

December 15, 1946 

CAR 3 Normal, Flash resistant or flame resistant if 

(1949-1965) Utility and smoking is allowed in a specific 

Effective Acrobatic compartment 

November 1, 1949 

Part 23 Normal, Flame resistant 

Through Utility and 

Amendment 23-22 Acrobatic 

(1965-1978) 

Effective 

December 5, 1978 

Part 23 (78-present) Normal, Flame resistant or selfextinguishing 

Amendment 23-23 Utility and per Appendix F for materials and 

Subs located Acrobatic on the cabin side of the firewall 

and Subsequent 

(1978-present) 

Effective 

December 1, 1978

154 
04/12/05 AC 23-17B 

Part 23 Commuter Self-extinguishing per Appendix F 

Amendment 23-34 except for small parts in 

and Subsequent § 23.853(d)(3)(v) 

(1987-present) 

Effective 

February 17, 1987 

Cargo and Baggage Compartment Fire Protection, § 23.855 

Part 23 Normal, Meet provisions of § 23.853(d)(3) 

Amendment 23-49 Utility and 

(1996-present) Acrobatic 

Effective 

March 11, 1996 

Part 23 Commuter Meet provisions of § 23.853(d)(3) 

Amendment 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 

Amendment 23-49 Utility, § 23.1182 and wire insulation self 

Effective Acrobatic and extinguishing per Appendix F, 

March 11, 1996 Commuter 60-degree angle test 

1. For 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 based on bill of materials and roll identification. The FAA 

Form 8110-3 should state it applies to the specific batch/roll/sheet for which the test 

was conducted. 

NOTE 16: 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 parts. These parts should be conformed per 

14 CFR, part 21, § 21.33, before 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 DERs have not always submitted FAA Forms 8110-3 with make and 

155 
AC 23-17B 04/12/05 

model information, FAA Forms 8110-3 must always be complete, including the make 

and model information.

156 
04/12/05 AC 23-17B 

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 build-up panel with an interior material, adhesive and composite 

structure, unless it is shown the interior material does not allow an ignition source to 

penetrate it. 

e. Interior flammability tests may be required with build-up samples. Experience 

has shown the thin exposed layer can burn away and expose the adhesive layer, 

which often is flammable and would contribute significantly to propagating 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). 

Amendment 23-14 and Subsequent 

Notice 71-13 explains this amendment level as follows: “The proposal makes 

provision for illuminated signs in airplanes having separate compartments. Because 

of increased complexity of systems carrying flammable fluids, the regulations must 

provide protection.” 

Amendment 23-23 and Subsequent 

Requirements for materials on the cabin side of the firewall were added because of 

this amendment. The following explanation in NPRM 75-26 states: “There have 

been instances where cabin upholstery has been ignited by engine fires even though 

the flame did not penetrate the firewall. This proposal would broaden Sec. 23.853 to 

require materials located on or adjacent to the cabin side of the firewall to be selfextinguishing or otherwise protected to prevent fire within the cabin due to engine 

fire.” 

Amendment 23-25 and Subsequent 

The proposed revision to NPRM 75-10 changed the “No Smoking” rules. The 

following explanation states: “The purpose of this proposal and the similar 

proposals to Secs. 25.853(c), 27.853(c) and 29.853(c) is to update and make 

consistent the certification requirements necessary to permit smoking in aircraft.”

157 
AC 23-17B 04/12/05 

Amendment 23-34 and Subsequent 

This change added requirements for commuter category airplanes. The explanation 

in NPRM 83-17 states the following: “SFAR 41, Sec. 7(b) and (c), and Secs. 91.56 

and 135.170 of the FAR.” 

158 
04/12/05 AC 23-17B 

23.855 Cargo and baggage compartment fire protection

The corresponding rule in CAR 3 is CAR 3.392. 

There is no corresponding rule in the Airship Design Criteria. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this rule as follows: 

“Proposed paragraph (a) would require all sources of heat within each cargo and 

baggage compartment that are capable of igniting the compartment contents to be 

shielded and insulated to prevent such ignition. 

Proposed paragraph (b) would require cargo and baggage compartments to be 

constructed of materials that meet the appropriate provisions of Sec. 23.853(d)(3). 

Currently these requirements apply to commuter category airplanes and to the 

materials used in the compartments of these airplanes. The proposed new 

requirement would expand this applicability to the cargo and baggage compartments 

of all part 23 airplanes. In effect, the proposed new requirement would require 

materials that are self-extinguishing rather than flame resistant as currently required 

under Sec. 23.787(d). 

Proposed new paragraph (c) would add new fire protection requirements for cargo 

and baggage compartments for commuter category airplanes. The proposed rule 

would require on of the following alternatives: (1) Either the compartment must be 

located where pilots seated at their duty station would easily discover the fire or the 

compartment must be equipped with a smoke or fire detector system to warn the 

pilot's station. The compartment must also provide access to the compartment with a 

fire extinguisher. (2) The compartment may be inaccessible, but must be equipped 

with a fire detector system that warns the pilot station, and the compartment must 

have ceiling and sidewall floor panels constructed of materials that have been 

subjected to and meet the vertical self-extinguishing tests of Appendix F of this part. 

(3) The compartment must be constructed and sealed to contain any fire. 

The proposed new section is necessary for several reasons. The proposals for 

additional requirements for commuter category airplane cargo and baggage 

compartments were developed after an examination of reported incidents of inflight 

fires and their causes. Although most of these incidents of inflight fires occurred on 

transport category airplanes, the reported sources of the fires showed that the fires 

originate from sources, such as matches in the pockets of clothing, that are as likely 

to be found on part 23 airplanes as on transport category airplanes. The same 

potential for inflight fires exists on commuter category airplanes and adequate 

protection should be provided.” 

159 
AC 23-17B 04/12/05 

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. 

160 
04/12/05 AC 23-17B 

23.859 Combustion heater fire protection

The corresponding rule in CAR 3 is CAR 3.388. 

There is no corresponding rule in the Airship Design Criteria. 

Amendment 23-5 and Subsequent 

A proposed revision to NPRM 65-43 explains this amendment as follows: 

“Secs. 23.859, 25.859, 27,859, and 29,859 would be amended to clarify, without 

changing, the conditions under which the "means independent of the components 

provided for the normal continuous control of combustion heater air temperature, 

airflow, and fuel flow" must be provided. The intent of this requirement is to require 

independent means of control when any one of the listed conditions exists.” 

Amendment 23-27 and Subsequent 

This rule was changed with the following explanation in Docket 20052: “The design 

complexity and operational altitude and weather capability of current Part 23 pistonpowered airplanes may necessitate installation of combustion heaters for occupant 

survival and airplane heating to counteract windshield icing conditions. Part 25 

presently contains updated type certification requirements for combustion heater fire 

protection that are applicable regardless of the size of airplane in which the heater is 

installed. Among these requirements are those for fireproof air ducts and protection 

of the ventilating airstream from backfire and reverse burning, as recommended by 

the NTSB for inclusion in Part 23. In addition, Part 25 designates fire zones, 

requires fire detection means, contains heater control requirements, sets forth air 

intake location standards, gives heater exhaust requirements, and provides for fuel 

system protection and drainage. These requirements constitute the necessary 

minimum safety standards for combustion heater fire protection for any highperformance airplane. 

Because of the increasing sophistication of normal category small airplanes and the 

expanded operating environment for which they are designed and used, the FAA has 

determined that more stringent type certification fire protection and detection 

standards are needed for combustion heaters when such heaters are installed in these 

airplanes. In view of the technical similarities between these airplanes and airplanes 

designed to Part 25 standards, the Part 25 standards for combustion heater fire 

protection are substantially adopted as type certification requirements for Part 23.” 

161 
AC 23-17B 04/12/05 

23.863 Flammable fluid fire protection 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.40.

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 

likelihood of ignition sources and leakage being present. This is followed by a means 

to minimize the effects once a fire has occurred by extinguishing, ventilation, 

isolation, and drainage. The rule does not go so far 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 likelihood 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 either 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: 

either (1) shrouding (sealing off) of all potential ignition sources, or (2) shrouding or 

sealing off all flammable fluid/vapor sources. In either case, it should be discovered 

the means would continue to serve its function following any single failure of the 

system or component it is isolating from the area. 

162 
04/12/05 AC 23-17B 

23.865 Fire protection of flight controls, engine mounts, and other flight structure 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.41.

Amendment 23-14 and Subsequent

A proposed revision to NPRM 71-13 explains this rule as follows: “Experience has 

indicated the need to ensure that flight controls and flight structure in powerplant fire 

zones remains operable in the event of a fire. The proposal requires this.” 

Amendment 23-45 and Subsequent

A proposed revision to NPRM 90-18 added “engine mounts” to the rule with the 

following explanation: “This proposal clarifies existing Section 23.865 by excluding 

those portions of the engine mount certificated with the engine from this section. 

Additionally, a clarification is provided to address the allowable damage expected on 

engine isolators.” 

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 on engine mounts is 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 enough 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 are 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 do 

not require testing, means of installing the shielding such as sealers and adhesives 

should be shown not to 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 show 

163 
AC 23-17B 04/12/05 

that it can withstand the loads expected to occur during completion of the flight. 

These loads can be treated as ultimate loads. Without a more rationale determination 

of the expected flight loads, a load factor of 70 percent of maneuver load or 40 

percent of gust load is superposed with the loads from the engine thrust and torque 

for maximum continuous power. After five minutes and until the end of 15 minutes 

the engine may be assumed to be shutdown. 

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.” 

The intent of this section in the rules is to require the materials and components, in 

the designated fire zone, that are essential to flight safety: 1) be fabricated either from

a material meeting defining fireproof under Title 14, part 1.1, or 2) be shown 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: 1) maintaining their integrity or 2) performing their function 

under the conditions of fire in the designated fire zone. 

The fire condition characterized by a 2000-degree Fahrenheit (F) flame can be treated 

as a failure condition that should not prevent continued safe flight and landing for at 

least 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 considered when 

establishing the remaining structural capacity. Freedom from flutter and whirl mode 

should also be shown. 

Compliance with this requirement must be shown by tests, or by analysis supported 

by tests. Assessing heat effects needs to include all heat transfer mechanisms that 

may occur. For composite structure, the long-term environmental effects that may 

degrade the mechanical properties of the 

164 
04/12/05 AC 23-17B 

structures also need to be considered. These may include the effects due to moisture 

and steam pressure. 

EASA AMC 23.865 is acceptable for FAA certification. 

165 
AC 23-17B 04/12/05 

ELECTRICAL BONDING AND LIGHTNING PROTECTION 

23.867 Electrical bonding and protection against lightning and static electricity 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.42.

Amendment 23-7 and Subsequent 

A proposed revision to NPRM 67-14 explains this rule as follows: “Small airplane 

operations involving potential exposure to lightning strikes are increasing. The 

hazardous effects of lightning on aircraft are well known. Economically practical 

means of protecting aircraft from such effects have been under study for several 

years. In order to prevent undue dictation of design, and yet, on the other hand, 

avoid an unnecessarily vague standard, this proposal summarizes the design 

objectives derived from the studies and experience obtained in the field of lightning 

protection of aircraft structures.” 

Lightning protection of Visual Flight Rules (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 relevant 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 spent 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 airplanes, an applicant should submit a certification test plan 

describing the analyses/testing to be used to show lightning protection effectiveness. 

At a minimum, the test plan should describe the sub-components, components, and 

systems to be tested and include: 

a. Zone definition of the entire aircraft (show how this will be accomplished—strike 

attachment model tests and similar design). 

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.

166 
04/12/05 AC 23-17B 

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 drawn. 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. 

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,” 

November 1, 1999. 

SAE ARP 5413, “Certification of Aircraft Electrical/Electronic Systems for the 

Indirect Effects of Lightning,” (being incorporated into an AC), May 14, 2002. 

SAE ARP 5414, “Aircraft Lightning Zoning,” December 1, 1999. 

SAE ARP 5415, “Users Manual for Certification of Aircraft Electrical/Electronic 

Systems for the Indirect Effects of Lightning,” May 14, 2002. 

SAE ARP 5577, "Aircraft Lightning Direct Effects Certification,” September 30, 

2002. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains the change in the title of § 23.867 as 

follows: “This proposed revision would change the heading that precedes the section 

from "Lightning Evaluation" to "Electrical Bonding and Lightning Protection." It 

would also revise the section heading form "Lightning protection of structures" to 

"Electrical bonding and protection against lightning and static electricity." The 

proposed revisions more accurately clarify the content of the section.” 

167 
AC 23-17B 04/12/05 

MISCELLANEOUS 

23.871 Leveling means

The corresponding rule in CAR 3 is CAR 3.401. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 4.50. 

Amendment 23-7 and Subsequent 

The original rule required “reference marks.” 

This amendment has the following explanation in NPRM 67-14: “any means, not 

just reference marks, are acceptable for leveling the airplane on the ground.” 

168 
04/12/05 AC 23-17B 

Subpart F—Equipment 

GENERAL 

23.1301 Function and installation 

The corresponding rules in CAR 3 are CAR 3.651 and 3.652. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.1.

Original Issue and Subsequent 

When performing 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 TSO requirements that are 

applicable for that instrument or equipment or equivalent requirements. 

Amendment 23-7 

A proposed revision to NPRM 67-14 explains this change as follows: “This proposal 

would remove the unnecessary distinction between radio and other equipment that 

now exists in Sec. 23.1301 (a) (2) and (3), and extend the provisions of Sec. 23.1431 

to all equipment, not only radio equipment. There is no basis for distinguishing 

between radio and other equipment so far as general functioning requirements are 

concerned. The other changes are intended to bring the sections in closer conformity 

with the corresponding sections in Part 25 through 29, without substantive change.”

169 
AC 23-17B 04/12/05 

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 under prescribed limitations. 

Therefore, if an instrument manufacturer specifies any allowable installation 

requirements (that is, panel slope for gyroscopic instruments), the installer should 

stay within the limitation(s). 

Each piece of installed equipment must be either labeled about its identification, 

function, 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 the 

installed equipment makes necessary. 

See ACs 23-8B, “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 the FAA 

will verify as part of the certification project. 

Section 23.1301 requires that instruments be installed under prescribed limitations. 

Therefore, if an instrument manufacturer specifies any allowable installation 

requirements (that is, panel slope for gyroscopic instruments), the installer should 

stay within the limitation. We recommend the slope be no more than 15 degrees. If 

applicants want a slope greater than 15 degrees, they should show decisively by tests 

or analyses the instrument will function properly when subjected to all expected 

airplane maneuvers. 

Each piece of installed equipment must be either labeled about its identification, 

function, 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 the 

installed equipment makes necessary. 

170 
04/12/05 AC 23-17B 

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 by 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 “SAOnly” 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 flown 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. 

Instruments and equipment installed to meet the requirements of § 91.205 should 

meet the TSO or equivalent requirements. TSO 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 is 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 (for example, VFR, Instrument Flight 

Rules (IFR), day or night) and the meteorological conditions (for example, 

icing) to which operating the airplane is limited, or from which it is 

restricted, must be established appropriate to the installed equipment. 

171 
AC 23-17B 04/12/05 

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 VFR or IFR. 14 CFR, part 91, § 91.205 requires for 

heading information; under VFR operation, a magnetic non-stabilized direction 

indicator (that is 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 TSOs and as acceptable means of compliance to § 23.1301 

in ACs, 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 

±10 degrees or have a correction card, placard, or a back-up gyroscopic direction 

indication provided indicator has an accuracy better than ±10 degrees. If the sole 

purpose of the gyroscopic direction indication is for backing up the magnetic nonstabilized direction indicator, then the accuracy of displayed headings can also be to 

±10 degrees. 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 

±4 degrees on the ground or ±6 degrees 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 because of: 

• the equipment itself,

• the current flow in any item of electrical equipment and its associated wiring, 

• the movement of any component, (for example controls or undercarriage), and

172 
04/12/05 AC 23-17B 

• the proximity of any item of equipment containing magnetic material direction 

indicators. 

COMPARATOR MONITOR 

Some systems installations 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: 

• six degrees in stabilized level flight. 

• six degrees plus 1⁄2 of the bank angle; or 

• 12 degrees with a bank angle greater than six degrees; 

• The alert function can be disabled at a bank angle greater than 20 degrees. 

• An alert is provided if the condition exceeds 60 seconds, but allow two minutes 

for a turn error as stated in the TSO. 

The six-degrees trip point during level flight allows a heading error of as much as 12 

degrees. This comprises one system at the six degrees in-flight tolerance limit while 

the other system, presumably with some malfunction, has an error of 12 degrees 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 either the TSO or equivalent requirements. Either the TSO or equivalent 

requirements are an acceptable means of compliance for either the instrument or 

equipment standards for installation in small airplanes. The relevant requirements in 

14 CFR, part 23, for the basis of this statement is 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 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. 

173 
AC 23-17B 04/12/05 

Section 23.1525, Kinds of operation.

The kinds of operation authorized (for example, 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 

for VFR operation by § 91.205. The required accuracy for magnetic non-stabilized 

direction indicator is prescribed in § 23.1327 as ±10 degrees 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 ±10 

degrees. 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 either 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 TSOs, 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 180 degrees in 

one minute at a true airspeed of 180 Miles Per Hour (MPH) (289 km/h) shall be 

within two degrees.” During stabilized and under environmental conditions, the test 

requirement for the error tolerance is two degrees. 

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 23-8B, “Flight Test Guide for Certification of Part 23 Airplanes," for 

additional guidance. 

174 
04/12/05 AC 23-17B 

23.1303 Flight and navigation instruments 

The corresponding rule in CAR 3 is CAR 3.655. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.2.

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 servo-corrected 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. Providing 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: 

1) identified by an appropriate part number, 2) marked clearly for use only on the 

airplane on which they are calibrated, and 3) give available information to the pilot to 

enable manual correction computations at airspeeds other than those used in 

designing the instrument correction. 

Amendment 23-17 and Subsequent

A proposed revision to NPRM 75-10 explains this amendment as follows: “The 

proposal would update Sec. 23.1303 to require a free air temperature indicator for 

turbine engine powered airplanes. The proposal would also update the section to 

provide for a speed warning device for turbine engine powered airplanes and certain 

other airplanes for which VMO/MMO and VD/MD have been established. The speed 

warning device proposal is similar to the requirement contained in Part 25 to the 

extent practicable for Part 23 airplanes. It should be noted that under the proposal 

turbopropeller powered airplanes would be required to have installed both a free air 

temperature indicator and speed warning device since service experience does not 

indicate any need to differentiate those airplanes from other turbine engine powered 

airplanes with respect to the need for such equipment.” 

175 
AC 23-17B 04/12/05 

See AC 23-8B, “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 

issued for the requirements of altimeter systems, only pneumatic altimeters were 

envisaged. 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 many instances 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. 

b. In assessing an electrically powered altimeter with pneumatic reversion capability, 

the means of providing continuous and usable altitude information should be 

considered on 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 instruments 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. Satisfactory 

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 pilots position provided a pneumatic altimeter is found on 

the instrument panel to be usable from any pilots' flight position. 

Radar Altimeters 

Radar altitude is not required for part 23, so it could be installed as no-hazard if it is a 

stand-alone installation of a TSO-C67, “Radar Altimeter,” with structural 

implications of the antenna considered, on a nonessential bus, with circuit protection, 

176 
04/12/05 AC 23-17B 

and separate annunciation from other aircraft warning systems. In this case, a field 

approval could be acceptable. 

If it is tied in to an approved warning system, or if it is part of a required TAWS, 

or if it is part of a precision landing system (Cat II or Cat III), then a STC is 

required. 

For TAWS, use AC 23-18 and TSO-C151. 

For precision landing, use 14 CFR, part 91, Appendix A. 

Guidance On 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 altitudeencoding altimeter: 

1. Install a monitor between another altitude source and the primary, encoded 

altimeter. If installed, a 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 

3. Install a TAWS. 

TAWS Class A and B are required for turbine-powered airplanes configured with 

six or more passenger seats operating under parts 91 and 135 and all turbinepowered airplanes operating under part 121. (See AC 23-18 for more details). 

We plan to initiate a rulemaking project for item 1. 

177 
AC 23-17B 04/12/05 

Long-Range Navigation-C System (LORAN-C) 

See AC 20-121A, “Airworthiness Approval of Loran-C Navigation Systems for use 

in the U.S. National & Airspace System (NAS) and Alaska,” for information on 

LORAN-C installations. 

Global Positioning System (GPS) 

See AC 20-138A, “Airworthiness Approval of Global Navigation Satellite System

(GNSS) Equipment”, December 22, 2003. 

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 on the intent of paragraph (a)(3)(x) of TSO-C129 as it applies 

to approach qualified (Class A1) GPS receivers. The paragraph states in part that the 

databases would contain all Standard Instrument Departures (SIDs) and all Standard 

Terminal Arrival Routes (STARs). 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 SIDs are not stored and selectable by name, then a limitation must be placed 

in the AFMS stating that SIDs must be entered before departure. 

- If all named waypoints of the STARs are stored individually in the database (not 

selectable as a procedure), then a limitation must be placed in the AFMS stating 

that GPS equipment is not approved for conducting STARs. This limitation may be 

removed as part of an installation approval for aircraft requiring more that one pilot 

crewmember. 

Terrain Awareness Warning System (TAWS) 

This guidance is in response to a request for clarification on follow-on installation 

field approval by the FSDO) for the TAWS, Class B, for part 23 airplanes. The 

questions regarded field approvals by the Airworthiness ASI approving an AFMS, or 

supplemental AFM. 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, “FollowOn Approval of Class B Terrain Awareness and Warning Systems (TAWS) 

178 
04/12/05 AC 23-17B 

(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), and 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 told that some ACOs have stopped 

certification projects of TAWS with electronic displays. The Small Airplane 

Directorate did not issue any policy to stop any installations of TAWS B in part 23 

airplanes. This policy is intended to allow current and future installations of TAWS

equipment, which is not only required for some part 23 airplanes, but also proven to 

increase safety. 

The Small Airplane Directorates position is 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 improvements: 1) improved 

situational awareness, and 2) fewer altimeter errors (for example, because of groundpilot communication/interpretation errors, pilot setting errors, static source errors, 

static system failures, icing effects, pressure and temperature changes, and nonstandard pressure gradients). However, as the current universal standard, pilots must 

use their primary barometric based altimeter instrument for navigation within the 

NAS. Thus, the MSL label on the TAWS altitude display, without clarification of its 

geometrically calculated multi-sensor source, could be misleading because of 

customary pilot expectations of a barometric based altimeter source. Nonetheless, 

now 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 (AD) for previous installations. The need or 

nature of a future labeling change is still being debated within the FAA. Until 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 according to the following 

paragraphs. 

1. The AFM or 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 on the use of the terrain awareness display. The 

terrain awareness display is intended to serve as a situational awareness tool only and 

179 
AC 23-17B 04/12/05 

may not provide the accuracy and/or fidelity on which to solely base terrain or 

obstacle avoidance maneuvering decisions. 

2. The AFM or 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 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 information to all installations until or 

unless they develop a software change that will re-label geometric altitude something 

else such as “GPSA.” 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 requires the direction indicator be a nonstabilized magnetic compass based on the following explanation: “This proposal 

would clarify the intent of the regulations for a magnetic nonstabilized direction 

indicator. The nonstablized magnetic direction indicator, which does not require 

power from the airplane's electrical systems, provides directional information to the 

pilot when all other directional navigation systems have failed due to loss of power. 

There have been contentions that a magnetic direction indicator with a remote 

magnetic sensor should be acceptable because only one magnetic direction indicator 

is required and it is more accurate for navigation. 

When the requirement for a magnetic direction indicator was promulgated, only a 

nonstabilized magnetic direction indicator (magnetic compass) was envisioned. 

Requirements in the operating rules for various kinds of operations are in addition to 

180 
04/12/05 AC 23-17B 

these basic requirements for type certification. The minimum level of safety 

established by these regulations was based on the reliability and failure modes of the 

nonstabilized magnetic compass and its ability to provide continuous heading 

information due to its functional independency.” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this amendment as follows: “The lead 

in for Sec. 23.1303(a) would be revised to clarify that the instruments required by this 

section are the minimum ones required. Also, Sec. 23.1303(d) would add a 

requirement for those airplanes whose performance must be based on weight, 

altitude, and temperature to be equipped with a free air temperature indicator. A 

new sentence added to Sec. 23.1303(e)(2) would state that nuisance overspeed 

warnings should not occur at lower speeds where pilots might ignore the warning. A 

new paragraph (f) would propose requirements for attitude instruments that include a 

means for flightcrew members to adjust the reference symbol. Finally, it would add a 

new paragraph (g) to define certain specific instruments required for a commuter 

category airplane. 

The proposal for Sec. 23.1303(e)(2) was developed following a Joint Aviation 

Authority recommendation that the warning should not occur below the maximum 

operating limit speed (VMO/MMO)/. To determine the effect that this recommended 

VMO/MMO limit would have on the design of overspeed warning devices, the FAA 

contacted several equipment manufactures. These manufacturers responded that it 

would be possible to establish a lower limit at VMO/MMO, but that the design 

changes needed to ensure that the warning occurred between the presently required 

upper limit and the recommended lower limit would be very expensive. 

The FAA notes that no known safety problem justifies that cost of these design 

changes. However, the FAA is also aware that if warnings of any type occur when 

the pilots know that no particular problem exists, such warnings may become a 

nuisance. If warnings become a nuisance, a pilot may disregard a warning when the 

airplane is approaching a flight speed where an unsafe flight condition may occur. 

Regulatory action is therefore needed to ensure that the warning will occur within 

appropriate speed limits. Proposed Sec. 23.1303(e)(2) would require manufacturers 

to establish a lower speed limit so that nuisance overspeed warnings will not occur. 

The manufacturer would be required to show that this limit is appropriate for the 

airplane design but would not be required to set this lower limit at one specific speed, 

such as VMO/MMO, which would be costly to achieve. 

A new Sec. 23.1303(f) is proposed because attitude instruments are available that 

provide a means accessible to the flightcrew members, for adjusting the reference 

symbol through ranges that could result in unsafe pitch angles in small airplanes. 

These instruments were developed for airplanes that use high pitch angles for 

181 
AC 23-17B 04/12/05 

approved climb or descent gradients. By permitting these airplanes to use 

instruments that can be adjusted for these higher pitch angles, pilots are able to 

maintain the design gradients using an instrument that provides a normal indication 

at that pitch. 

If each attitude instruments are installed in small airplanes, pilots could adjust the 

reference symbol to ranges that could result in unsafe pitch angles. The 

recommendation showed that some instruments can be adjusted to result in pitch 

angles that are nearly the same as the pitch angle that many small airplanes achieve 

before stalling. To preclude potential cases of unwanted pitch adjustments of attitude 

instruments installed in small airplanes, Sec. 23.1303(f) proposes to limit the 

adjustment range to that limit that is needed for parallax correction. 

Proposed new Sec. 23.1303(g) would identify specific instruments, and limits of those 

instruments, required for commuter category airplanes. When the JAA initiated their 

consideration of commuter category airplanes, one of the proposals they received 

recommended adding the instrument requirements of Sec. 25.1303 to part 23 for 

commuter category airplanes. In considering this recommendation, a review of the 

requirements showed that many instruments required under Sec. 25.1303 are 

presently required by the operating rules. In addition, Sec. 23.1583(h) requires a list 

of the equipment that must be installed for the kinds of operation for which the 

airplane is approved. Based on the review, it was determined that many of the 

requirements in Sec. 25.1303 would be redundant, and the recommendation was not 

accepted. 

In considering a portion of the recommendation to require a third attitude 

instrument, the FAA noted that Sec. 91.531(a)(3) requires a commuter category 

airplane of ten or more passengers to be operated with a second-in-command and 

that Sec. 23.1321 requires flight and navigation instruments for such required pilot. 

Accordingly, two attitude instruments are required for a ten passenger, IFR approved 

commuter category airplane. Service experience has shown that failures of an 

attitude instrument system can occur where there will be a time period in which the 

indicator appears to be working but is providing incorrect information. During such 

a failure of one instrument in an airplane equipped with only two instruments, the 

pilots may have difficulty determining which instrument to follow, and hazardous 

flight attitudes may result. A third attitude instrument would allow the crew to retain 

reliable attitude information at all times, and thus the proposed rule would require a 

third attitude instrument for commuter airplanes operated by two pilots.” 

182 
04/12/05 AC 23-17B 

23.1305 Powerplant instruments 

The corresponding rules in CAR 3 are CAR 3.655 and 3.675. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.3.

Original Issue and Subsequent

a. Fuel Pressure Indication 

Paragraph (b)(4)(ii) of § 23.1305 requires fuel pressure indicators for pump-fed 

engines. An ELOS 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 either a placard or 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.” In consideration of the 

policy in item 6d of AC 20-88A, where the rate of change is small or nearly 

steady state (that is, 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 installing 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 airplanes for 

which those regulations are applicable. Listed below are the applicable part 23 

sections with the related CAR sections shown in parentheses: 

183 
AC 23-17B 04/12/05 

Part 23 Sections 

§ 23.773 (3.382) 

§ 23.955 (3.433) 

§ 23.961 (3.438) 

§ 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. 

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. The accuracy of these readings is dependent on 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 Amendment 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 on 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.

184 
04/12/05 AC 23-17B 

3. DISCUSSION 

a. Fuel Pressure and Fuel Quantity Indicator 

(1) A fuel pressure indicator is required for pump-fed engines according to 

§ 23.1305(g). It is intended to monitor metered fuel pressure at the inlet 

to the 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 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 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 allow 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 because of temperature 

variations, malfunctions in the fuel system such as leaks, siphoning 

actions, collapsed bladders, and other factors. So, total fuel remaining 

should be verified with the fuel quantity indicator. According to 

§ 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. Therefore, 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 

185 
AC 23-17B 04/12/05 

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 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 according to § 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 if 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 for 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 TC, 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 flight-engine-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 enginedriven 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 

186 
04/12/05 AC 23-17B 

§ 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 

according to § 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 and 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 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 ELOS for the airplane shows 

that replacing the fuel pressure indicator with a flowmeter 

will still meet the applicable airworthiness requirement. A 

finding of ELOS should confirm the instrumentation 

provided by the fuel flowmeter is satisfactory, reliable, 

and safe under all reasonably foreseeable operating 

conditions.

187 
AC 23-17B 04/12/05 

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 TC or STC for either the 

airframe or the engine. FAA approval is obtained after the applicant shows 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 allows a digital flowmeter with 

specific instructions. An improper installation not only will risk the safety of the 

present designs, but could also increase the likelihood of system failure. 

Installations should comply with the airworthiness regulations and with the 

manufacturers installation criteria. 

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 

pressure-type flowmeter. 

(2) Fuel Injection System with a Constant Pressure Pump 

Installation of a digital or analog fuel flowmeter may replace the analog 

pressure-type 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 ELOS. 

(3) General Considerations 

Changes to the fuel systems should be evaluated for: 1) fuel flow rates, 2) 

maximum allowable pressure drop, 3) hot weather operations, 4) vibration 

and loads on lines and fittings, 5) fire protection, and 6) powerplant 

instruments, including effects of glare and reflections on instruments in the 

pilot compartment. An engineering analysis should be made to assure: 1) 

good engineering practices are incorporated in the design and 2) the 

installation is in accordance with airworthiness standards of §§ 23.773, 

23.955, 23.961, 23.993, 23.1183, 23.1191, 23.1301, 23.1309, 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 on each particular installation.

188 
04/12/05 AC 23-17B 

c. Evaluation 

Modification of the approved fuel system may have major effects; therefore, an 

evaluation should be conducted to confirm 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; therefore, 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. 

(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 discover 

their capability to supply the required fuel flow rate at the higher pressure. 

Flight tests for turbocharged engines may be required to determine the 

minimum fuel injector inlet pressure meets the engine TCDS 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 provided with the safe operating limits. 

A placard should be 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) 

Either a FMS or AFMS, or placards, if appropriate, should be prepared by the 

applicant. The information should be presented for FAA approval in the 

following sections: 

189 
AC 23-17B 04/12/05 

(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 

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 on 

the initial fuel supply programmed into the computer before the start of 

each flight. Doubts about initial fuel supply and total fuel remaining can be 

because of an uneven ramp, unusual loading, volume changes of the fuel 

because of 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 the pilot determine the fuel programmed into the computer is 

the same as the usable fuel on board the airplane. 

(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 with placards. 

Amendment 23-7 and Subsequent 

A proposed revision to NPRM 67-14 explains this amendment that added turbine 

engine requirements as follows: “The powerplant instruments in present Sec. 

23.1305 are not appropriate for turbine engine powered airplanes. The instruments 

in this proposal are those that experience has shown to be necessary for the safe 

operation of those airplanes.” 

Changes to the NPRM explain Final Rule, Docket 8083 as follows: “The notice 

proposed to amend Sec. 23.1305 to require specified powerplant instruments for 

turbine engine powered aircraft. One comment objected to the requirement for a 

propeller blade angle sensor. However, the FAA considers that the reliability of 

direct sensing is essential, considering the critical nature of an unwanted travel of 

the blade below the flight low pitch position. In response to another comment, the 

FAA has withdrawn the proposed requirement for a reverse-blade angle sensor. The 

FAA is aware that such a sensor is not currently practical. The proposed 

requirement for firewarning indicator is also withdrawn consistent with the 

withdrawal of proposed new Sec. 23.1203. The notice also proposed to add a new 

Sec. 23.1307(c) to require, for turbosupercharger installations, specific instruments 

to allow monitoring of these installations. In response to a comment received, the 

proposed amendment has been 

190 
04/12/05 AC 23-17B 

changed to require these instruments only if limitations are established for either 

carburetor air inlet temperature or exhaust gas temperature and those limitations 

can be exceeded in operation, and the amendment has been adopted as Sec. 

23.1305(p). In addition, the cylinder head temperature indicator requirements of 

present Sec. 23.1337(e) are transferred to new Sec. 23.1305(f) inasmuch as it is a 

required instrument and should properly be included with the other required 

powerplant instruments. No substantive change is intended. In addition, the 

proposal concerning thrust indicators has been changed to specify the equivalent 

means, and the reference to an equivalent to a torque indicator has been deleted 

since there is no equivalent, and the purpose of the tachometer for turbine engines 

has been specified. Finally, the format of present Sec. 23.1305 has been changed to 

simply list the required instruments inasmuch as the requirement for each instrument 

is now set forth as to each instrument.” 

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 ELOS. 

4. Fuel Strainer or Filter Indicators 

Acceptable means of compliance for fuel strainers or filter indicators for turbineengine 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 

191 
AC 23-17B 04/12/05 

should be capable of sustained operation while operating with water in the 

fuel

192 
04/12/05 AC 23-17B 

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 an airframe-mounted 

filter is not installed, care should be taken to assure 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 so it can be readily 

inspected for operation before 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), because the rule 

requires an indicator, which is a cockpit-mounted indication. 

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. 

Diesel engines were not considered when this rule was written. For policy on diesel 

engine installations in small airplanes, see PS-ACE100-2002-004, Final Policy 

Statement; “Diesel Engine Installation.” 

Amendment 23-14 and Subsequent 

A proposed revision to NPRM 71-13 added requirements for: 1) an oil pressure 

warning means for each turbine engine and 2) an induction system air temperature 

indicator for each engine equipped with a preheater and having induction air 

temperature limitations which can be exceeded with preheat. No explanation was 

given. Amendment 23-14 is explained by Final Rule, Docket 11011, as follows: 

“Proposed Sec. 23.1305 (a), now Sec. 23.1305(q), has been revised to clarify that the 

only oil pressure warning means required is for low oil pressure.” 

Amendment 23-15 and Subsequent 

This amendment added these instruments: 

(q) For each turbine engine, an indicator to indicate the functioning of the 

powerplant ice protection system. 

193 
AC 23-17B 04/12/05 

(r) For each turbine engine, an indicator for each fuel strainer or filter to indicate the 

degree of contamination of the strainer or filter to the degree established as an 

operating limitation for the engine. 

(s) For each turbine engine, a warning means for each oil strainer or filter that has no 

bypass, to warn the pilot of the occurrence of contamination of the strainer or 

filter before attainment of contamination of the degree established as an operating 

limitation for the engine. 

(t) An indicator to indicate the functioning of any heater used to prevent ice clogging 

of fuel system components. 

A proposed revision to NPRM 71-12 explains Amendment 23-15 as follows: “The 

provisions of paragraph (q) would be transferred from present Sec. 23.1093. The 

indicator and the means proposed in paragraphs (r) and (s) are necessary to 

implement the requirements of proposed Secs. 23.997 and 23.1019, respectively. The 

indicator proposed in paragraph (t) is necessary in the event that a mechanical 

means, such as a fuel heater, is used in the airplane to heat the fuel as it passes 

through a filter or screen in the fuel system. Such an indicator could be used in 

meeting the requirements of proposed Sec. 23.951.” 

Amendment 23-15 is further explained by Final Rule, Docket 11010 as follows: “As 

a result of the issuance of Amendment 23-14 (38 FR 31816), which contains new 

paragraphs, Sec. 23.1305(q) and (r), the proposed paragraphs (q), (r), (s), and (t) are 

redesignated (s), (t), (u), and (v), respectively. The following discussion is keyed to 

the new designations. 

One commentator questioned whether the fuel strainer or filter indicator referred to 

in Sec. 23.1305(t) were required on all filters, even "last chance" filters. Consistent 

with the requirements applicable to the strainers or filters themselves, Sec. 23.1305(t)

has been revised to make clear that the indicator required is for a fuel strainer or 

filter required under Sec. 23.997. Similarly, Sec. 23.1305(t) and (u) have been 

reworded in order to be consistent with Secs. 23.997 and 23.1019, respectively, in 

regard to the degree of contamination that must be indicated. In response to a 

further comment, Sec. 23.1305(t) has been reworded to clarify that the desired 

indication is of the occurrence of contamination rather than the more stringent 

requirements of the degree of contamination as suggested in the notice. This change 

achieves consistency between paragraphs (t) and (u).

One commentator questioned whether other presently installed gauges for other 

functions could be used as "indicators" to indicate the functioning of a heater as 

required in paragraph (v). As discussed above, in connection with the indicators 

required for oil strainers or filters, the FAA anticipates that the requirement will be 

met by installation of gauges to indicate the functioning of heaters. However, if a 

clear and positive indication can be obtained from other gauges used to portray 

194 
04/12/05 AC 23-17B 

functions different than direct heater functioning, the requirements of the section are 

met.” 

Amendment 23-18 

A proposed revision to this amendment added a requirement for a fire-warning 

indicator for those airplanes required to comply with § 23.1203. 

Amendment 23-26 

A proposed revision to NPRM 75-31explains this amendment as follows: “The 

proposal would permit up to an eight-degree movement of the propeller blade below 

the flight low pitch position before an indication of the movement is required for the 

flight crew. This would be consistent with the corresponding requirements in Part 25 

and with the proposal to revise Sec. 35.23. The FAA does not believe it necessary to 

require immediate indication when the propeller blade moves past the flight low pitch 

stop.” 

Amendment 23-34 

A proposed revision to NPRM 83-17 added the following instruments: 

(f) A cylinder head temperature indicator for - 

(1) Each air-cooled engine with cowl flaps, and for each airplane for which 

compliance with Sec. 23.1041 is shown at a speed higher than VY; and 

(2) Each reciprocating engine-powered commuter category airplane. 

(h) A manifold pressure indicator for - 

(1) Each altitude engine; and 

(2) Each reciprocating engine-powered commuter category airplane. 

(k) A fuel flowmeter for - 

(1) Each turbine engine or fuel tank, if pilot action is required to maintain fuel flow 

within limits; and 

(2) Each turbine-powered commuter category airplane. 

“The source for this change was: part 135, Appendix A., § 58.”

195 
AC 23-17B 04/12/05 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 changed Amendment 23-43 as follows: “This 

proposal clarifies the powerplant instrument requirements by separating the 

reciprocating engine, turbine engine, turbojet/turbofan, and turbopropeller 

requirements; and adding requirements for coolant temperature indicator, fuel lowlevel warning, manifold pressure indicator, and chip detector indicator.” 

This amendment was added by Final Rule, Docket 26344 as follows: “This proposes 

to amend Sec. 23.1305 to clarify the powerplant instrument requirements by 

reorganizing the section and by defining the additional instruments that are required 

for the particular type of engine that is installed. Two commenters submitted 

comments on various paragraphs of this proposal. 

One commenter suggests that the proposal for Sec. 23.1305(a)(1) cross-reference 

Sec. 23.1337(b)(5). The FAA does not agree that the suggested Sec. 23.1337(b)(5) is 

the proper reference but does agree with the intent of this suggestion. In reviewing 

this comment, it is noted that the intent of the words, "or for each assembly of 

interconnected tanks that function as one tank" in proposed Sec. 23.1305(a)(1) is 

provided by current Sec. 23.1337(b)(4), and that other provisions of Sec. 23.1337(b)

address other applicable fuel quantity indicator requirements, such as their marking. 

To clarify the fuel quantity indicator requirement of part 23, the above quoted words 

from the proposal for Sec. 23.1305(a)(1) are removed and replaced by the words 

"installed in accordance with Sec. 23.1337(b). By this change, Sec. 23.1305 will 

make it clear that a fuel indicator is required for each tank and Sec. 23.1337 will 

provide the installation requirements for those indicators. 

Both commenters requested that the FAA make it clear that a dipstick is an 

acceptable oil quantity measuring device for meeting the requirement of proposed 

Sec. 23.1305(a)(4). The FAA notes that Sec. 23.1337(d) identifies an acceptable 

means of measuring the oil quantity and identifies a stick gauge as being one 

acceptable means. To provide the clarification requested by these commenters, the 

proposal for Sec. 23.1305(a)(4) is revised by adding the words, "which meets the 

requirements of Sec. 23.1337(d)" to the end of the proposal. 

One commenter asks the FAA to make it clear that N1 is an acceptable parameter 

that can be related to the thrust indication required by proposed Sec. 23.1305(d)(1). 

The FAA has reviewed this request and finds that the indication of the N1 speed is an 

acceptable means. For some installations, however, the applicant may be required to 

demonstrate that N1 is acceptable. By the discussion of this comment, the requested 

clarification is provided and proposed Sec. 23.1305(d)(1) is not revised.” 

196 
04/12/05 AC 23-17B 

Amendment 23-51 and Subsequent 

Amendment 23-51 was changed by a revision to NPRM 94-19 as follows: “The 

proposed revision to paragraph (b)(3) would delete paragraph (b)(3)(ii) which refers 

to compliance with Sec. 23.1041.” 

Amendment 23-52 and Subsequent 

A proposed revision to NPRM 94-37 explains this amendment as follows: “The 

intent of the fuel pressure indicator requirement for pump-fed engines is to advise the 

pilot of a fuel pressure deficiency before total engine failure. The term "indicator" in 

Sec. 23.1305 (b)(4) implies that the fuel pressure be constantly displayed. 

The proposal would change the current requirements in that a fuel pressure indicator 

or a fuel flow indicator would be acceptable. The fuel flow indicator would 

constantly display information that the pilot could use to evaluate engine power, fuel 

mixture, and other engine performance factors. Furthermore, it is technologically 

possible to have a microprocessor that monitors engine operation and triggers a 

warning if the fuel system operation does not match the other monitored engine 

trends. Therefore, this proposal would also change the rule to accept a means that 

monitors the fuel system and warns the pilot of any trend that could lead to engine 

failure. 

Accordingly, this proposal would adopt a performance standard, instead of a 

requirement for specific equipment. In this way, the designer could show compliance 

with paragraph (b) of the proposed by developing any design that monitors the fuel 

system and warns the pilot of any trend that could lead to engine failure. The 

Aviation Rulemaking Advisory Committee (ARAC) did not believe this would reduce 

the level of safety originally intended by the requirement. A warning light system 

could possibly alert the pilot sooner than if the pilot relied on an instrument panel 

scan to notice a trend in the fuel pressure indication. 

Microprocessing units that monitor engine operation and warn of fuel system 

problems have already been incorporated in transport aircraft and automobiles. 

Furthermore, pilots are not monitoring gauges like they use to; instead, they are 

increasingly relying on warnings to alert them. Late model automobiles, computers 

and other equipment are designed to protect the operators from mistakes by using 

built-in warnings. It is important to note that this NPRM does not propose to allow 

"idiot lights" to replace fuel pressure gauges. A light that comes on at the same time 

that the engine quits is useless. A warning light system that would comply with this 

proposal would be sophisticated enough to read transients and trends, and would 

give a useful warning to the pilot. The FAA expects this proposal to result in fuel 

systems that provide the pilot with useful engine operating information; thereby, it 

would offer more value to the operator. 

197 
AC 23-17B 04/12/05 

Today, fuel pumps are more reliable than those built in the 1940s and 50s. 

Consequently, airplane operators are more concerned about reducing engineoperating costs than they are about the probability of a fuel pump failure. 

A fuel flow indicator offers additional value compared to a fuel pressure indicator. It 

enables the operator to monitor the engine's fuel consumption and compare it to fuel 

consumption listed in the airplane flight manual. If a fuel monitoring system is 

installed that automatically controls the engine or helps the pilot to properly lean the 

fuel mixture, then engine operation would be optimized and the direct operating costs 

would go down through reduced fuel consumption. Reciprocating engines run better 

if the fuel to air mixture is leaned out according to the optimum (manufacturer's)

specified setting. Furthermore, fuel flow also rotates to power, and pilots can use 

fuel flow readings to quickly access the health of their engine during critical phases 

of flight, such as takeoff. 

Comprehensive engine monitors and redesigned electronic engine instrument 

displays are also being used in experimental aircraft. The FAA should encourage 

airplane manufacturers to utilize new technology to improve operation and reduce 

operating costs. New engine monitoring systems may improve reliability and engine 

life resulting in increased safety. 

The proposal would achieve the same safety objective as the current rule; the crew 

would have sufficient warning of any negative trend that could lead to partial or total 

engine failure. However, the proposal recognizes that this objective can be achieved 

by measuring fuel pressure, fuel flow, or with a "smart" fuel monitoring system.” 

This proposal was further explained by Final Rule, Docket 28011 as follows: 

“Transport Canada questions the ability to show compliance with the requirement in 

Sec. 23,1549 to identify maximum and, if applicable, minimum safe operating limits 

as well as the normal operating range of the instrument. This commenter points out 

that the typical fuel flow meter is a digital type, and it would be difficult for the 

applicant to provide equivalent markings. Engine manufacturers provide the 

information required by Sec. 23.1549, which is then usually transcribed to the 

installed fuel pressure gauge. It appears that this information would not be presented 

through the use of typical digital fuel flow meters. The commenter offers the 

following suggestion: "FAR 23.1549 was written with a traditional dial instrument in 

mind where the engine limitations could be easily displayed on the face of the unit 

and monitored by the crew. To allow flow meters or other fuel system monitors to 

satisfy the requirements of Sec. 23.1549 where such a gauge no longer exists, 

compliance could be shown by (1) different colors to indicate changing trends in 

system performance (e.g., amber color for a low pressure/flow condition, red for 

impending engine failure), or (2) placarding, if appropriate, to indicate the normal 

and abnormal operating ranges". 

198 
04/12/05 AC 23-17B 

The FAA agrees with the commenters' suggestions as an acceptable means of 

compliance with § 23.1549. Suggested items (1) and (2) above offer the pilot a 

means to determine fuel flow limitations, which may be needed if a fuel flow meter is 

installed. 

A commenter from Australia supports the proposal; however, the commenter feels the 

proposed text would require a monitoring system that provides a warning of any trend 

that could lead to engine failure, which is a difficult compliance requirement. The 

commenter further states: "The historic requirement, and the NPRM preamble, 

clearly addresses fuel pressure (as an indication of the availability of fuel flow) or 

fuel flow only. Such wording may stifle the development of monitoring 

instrumentation for small airplanes". The commenter suggests that, for clarification, 

the proposed text for Sec. 23.1305(b)(4)(ii), be amended to read as follows: "That 

continuously monitors the fuel system and warns the pilot of any fuel flow trend that 

could lead to engine failure". 

The FAA agrees with the commenter the proposed wording may be too broad, 

making compliance difficult or the system unnecessarily complex. The FAA 

encourages "smart" systems; however, the intent of the proposal was to warn the pilot 

of any fuel flow trend and, for that reason, the final rule and the preamble adopt the 

commenters' language. 

Section 23.1305 is adopted with the change in paragraph (b)(4)(ii) to add the words 

"fuel flow" before the word "trend".” 

199 
AC 23-17B 04/12/05 

23.1307 Miscellaneous equipment 

The corresponding rule in CAR 3 is CAR 3.655. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.5. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains a revision to this amendment as 

follows: “This proposal deletes the word "an approved" in Section 23.1307, and 

adds a new requirement that all equipment items necessary for the airplane to 

operate in the NAS to its maximum approved altitude and in all kinds of operations 

for which it is approved must be included in the type design. 

Conference proposal 432 recommended deleting Section 23.1307(a) in its entirety, 

because it was redundant to Section 23.785(b). Conference proposal 433 

recommended deleting only the word "approved" in Section 23.1307(a). 

Both Sections 23.1307 and 23.785(b) contain seat requirements; however, Section 

23.1307(a) only requires a seat or berth for each occupant, while Section 23.785(b)

contains specific design requirements for each seat, berth, safety belt, and shoulder 

harness, but does not include requirements for a seat or berth to be provided for each 

occupant. The requirements of Section 23.1307(a) were added by amendment 23-23 

to eliminate questions as to the maximum seating capacity and compliance with the 

emergency exit requirements. The FAA does not consider these requirements to be 

redundant and plans no further action on these recommendations. 

A new paragraph (c) is added because the FAA considers it necessary to clarify the 

type design requirements for part 23 airplanes relative to equipment items. 

Frequently, manufacturers have requested their airplanes be approved relative to 

structural, performance, and propulsion requirements for a specific altitude without 

also requesting approval of necessary equipment to operate at that altitude. The FAA 

considers it necessary for the type design to include all equipment necessary for the 

type design to include all equipment necessary for operation in accordance with the 

limitations required by Sections 23.1559 and 23.1583.” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21explains this rule as follows: “This proposal 

would remove the requirement of Sec. 23.1307(a) which is being added to Sec. 

23.785. The discussion of sec. 23.785 covers this change. Also, the provisions of

200 
04/12/05 AC 23-17B 

Sec. 23.1307(b)(1), (b)(2), and (b)(3), are being removed from Sec. 23.1307. These 

requirements have been previously added to Secs. 23.1361, 23.1351, and 23.1357, 

respectively; therefore, they are redundant and may be removed. The designator for 

paragraph (c) has also been removed from the remaining text of this section.” 

201 
AC 23-17B 04/12/05 

23.1309 Equipment, systems, and installations 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.6.

Amendment 23-14 and Subsequent

A proposed revision to NPRM 71-13 explains the adoption of this rule as follows: 

“Because of the increasing reliance on systems and equipment in the modern, 

complex, high performance airplanes, the FAA believes that the proposed 

amendments to Secs. 23.1301, 23.1305, and 23.1309 are necessary to insure the 

reliability of such systems and equipment.” 

Adoption of this rule is also explained by Final Rule, Docket 11011 as follows: “The 

other commentator objected to the proposed section on the ground that it could be 

interpreted to require the complete elimination of radio interference which is 

impossible. That is not the intent and the FAA does not agree with that 

interpretation. Similarly worded requirements already exist in the Federal Aviation 

Regulations and have not been misinterpreted to place an undue burden on 

manufacturers. However, based on further review, the proposed section has been 

revised to require that equipment installed in single engine airplanes be designed to 

minimize hazards to the airplane in the event of a probable malfunction or failure.” 

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” (that is, the functioning of the ice 

protection system). 

Software 

Order 8110.49 "Software Approval Guidelines", issued June 3, 2003, provides 

guidelines for the approval of software changes in legacy systems using 

RTCA DO-178B. 

See AC 23.1309-1C, “Equipment, Systems, and Installations in Part 23 Airplanes,” 

for additional guidance. 

202 
04/12/05 AC 23-17B 

Airplane Parachute Recovery Systems

Use of a parachute recovery system is not covered by part 23. Applications for 

installing this system will require special conditions.

Amendment 23-17 

A proposed revision to NPRM 75-10 had a minor editorial change in adding commas 

in the title. 

Amendments 23-34 through 23-40 

Amendment 23-34 added requirements for commuter category airplanes that were 

later adopted for all part 23 airplanes. A proposed revision to NPRM 83-17 “added 

the requirements from part 135, Appendix A., § 59”. 

Amendment 23-41 and Subsequent 

A proposed revision to NPRM 89-6 explains this amendment as follows: “Current 

Part 23 airplane airworthiness requirements are based on single-fault or fail-safe 

concepts and, when they were promulgated, the FAA did not envision use of complex, 

safety-critical systems in such airplanes. This proposal will require examination of 

systems and equipment for their criticality for continued safe flight and will permit 

the continued use of the existing reliability requirements of Part 23 for airplanes 

whose systems are not complex and do not perform safety-critical functions. For 

those cases where the manufacturer finds it necessary or desirable to include 

complex, safety-critical systems, the proposal includes requirements for identifying 

those systems and defines additional requirements needed for their certification. The 

proposed changes to Section 23.1309 are summarized as follows: 

Proposed paragraphs (a), (a)(1), (a)(2), and (a)(3) of Section 1309 are derived from 

current paragraphs (a), (b), and (c) of Section 23.1309. Under proposed Section 

23.1309(a)(1), the systems and installations, as well as the items of equipment, will be 

required to meet those requirements contained in current Section 23.1309(a). Under 

proposed Section 23.1309(a)(2), the systems and installations, as well as the items of 

equipment for single-engine airplanes will be required to meet those requirements 

contained in current Section 23.1309(c) and proposed Section 23.1309(a)(3) will 

require these items on a multiengine airplane to meet the requirements contained in 

current Section 23.1309(b). 

(2) A new Section 1309(b) is proposed which will require a detailed examination of 

each item of equipment, system, and installation. This examination is to determine 

whether a failure would effect the airplane's continued safe flight and landing. Each 

item of equipment, each system, and each installation identified by such an 

examination as being critical to the safe operation of the airplane would be required 

to meet additional requirements. This will permit the approval of more 

203 
AC 23-17B 04/12/05 

advanced systems, that were not envisioned when Section 23.1309 was added to Part 

23, without the need for special conditions. 

A new Section 23.1309(c) is proposed to require identification of loads which are 

"essential loads" and requires the airplane power sources for these loads to meet 

requirements consistent with other airplane airworthiness requirements. These 

requirements are substantially equivalent to section 59(b) of Appendix A to Part 135. 

A new Section 23.1309(d) is proposed to allow reduction of power loads when 

showing compliance with proposed Section 23.1309(c)(2). This provision is 

substantially equivalent to section 59(c)(3) of Appendix A to Part 135. 

A new Section 23.1309(e) is proposed to require that the design of each electrical 

system, each item of equipment, and each installation take into account critical 

environmental conditions. Critical environmental and atmospheric conditions would 

include radio frequency energy and direct and indirect effects of lightning. Section 

23.867 now requires the airplane structure to be protected from the effects of 

lightning and Section 23.954 now requires the airplane fuel system to be protected 

from the effects of lightning. 

A new Section 23.1309(f) is proposed to provide a definition of the systems to which 

the requirements of this section are applicable and specifically identifies certain 

items to which they are not applicable. 

Early in the development of minimum requirements for transport category airplanes, 

it was realized that more complex systems were being added to airplane designs and 

there was a need to include requirements which specifically addressed equipment, 

systems, and their installation. This was accomplished by the addition of Section 

4b.606 to Part 4b of the Civil Air Regulations (CAR), Amendment 4b-6, effective 

March 5, 1952, in part, accomplished this objective. The preamble to that 

amendment included the following in regard to that addition: 

An amendment clarifying the requirements for equipment, systems, and installations 

with regard to functioning and reliability is made in Subpart F. In addition, it 

specifies dual power supply for those installations the functioning of which is 

necessary to show compliance with the Civil Air Regulations. 

This requirement, as adopted by that amendment, was retained in CAR Part 4b until 

that regulation was recodified to the Federal Aviation Regulations, Part 25, effective 

February 1, 1965. At that time, this requirement, substantially unchanged, was 

identified as Section 25.1309. For many years the "single fault" or "fail safe" 

concept of this requirement, along with experience based on service-proven designs 

and engineering judgment, were used to successfully evaluate most airplane systems 

and equipment. 

204 
04/12/05 AC 23-17B 

However, in the late 1960's there appeared a number of safety-critical systems that 

were utilizing new technical complexity to accomplish safety-critical functions. Due 

to the increasing complexity of the technology being used to develop these systems, it 

was becoming increasingly difficult to apply engineering judgment as the only means 

of determining the effects or likelihood of failure conditions. The increasing difficulty 

in evaluating these complex systems, along with the potential hazards to the airplane 

that could result from their failure made it necessary to provide duplicate and 

triplicate systems to assure an acceptable level of safety. 

At about this same time, the development of rational methods for safety assessment of 

systems led to the conclusion that an inverse relationship should exist between the 

probability of a failure condition and its effect on the airplane. That is, the more 

serious the effect, the lower the probability must be that it will occur. 

The availability of this rational method for safety analysis of systems, along with the 

increasing difficulty in applying the then existing "single fault" or "fail safe" concept, 

prompted the FAA to propose an amendment to Section 25.1309 which would permit 

the use of this rational method as an acceptable means of accomplishing safety 

analysis. That proposed revision to Section 25.1309, which specified a level of safety 

in qualitative terms, and required assessments to be made, was adopted by 

Amendment 25-23 on April 1, 1970. 

At that time, the FAA also realized that more complex systems were being utilized in 

small airplanes and that there was a need to add a reliability requirement to Part 23. 

Accordingly, rulemaking action was initiated which resulted in Section 23.1309 being 

added to Part 23 by Amendment 23.14, effective December 20, 1973. The 

requirements of that new Section 23.1309 were similar to those of Section 25.1309 

prior to Amendment 25-23 which were based on the "single fault" or "fail safe" 

concept. At that time, it was not envisioned that complex safety-critical systems, 

utilizing technology now available, would be used in the designs of small airplanes. 

Therefore, there was no identified need to include provisions for use of the rational 

method of analysis in Part 23, as had been done in Part 25 by Amendment 25-23. 

Experience has shown that the envisioned rate of technical growth of systems used in 

small airplanes was inaccurate and that safety-critical systems are now being 

proposed for use on Part 23 airplanes. As with the earlier experience with Part 25, 

the FAA is finding that it is difficult to apply "single fault" concepts to these complex 

systems and to utilize the application of engineering judgment as the only means of 

determining the effects or likelihood of certain failure conditions. Accordingly, there 

is now a need to revise the reliability requirements of Part 23 to allow the use of the 

latest available rational method for safety analysis of these complex safety-critical 

systems to assure continuation of the level of safety intended for airplanes 

certificated to Part 23. Such safety critical systems are currently being proposed for 

approval and there is an urgent need to accomplish this proposed rule change. 

205 
AC 23-17B 04/12/05 

A recommendation for a complete change of Section 23.1309 was submitted as a part 

of the Small Airplane Airworthiness Review Program. That recommendation, 

conference proposal number 434, was discussed during the October 22-26, 1984, 

Small Airplane Airworthiness Review Program conference held in St. Louis, 

Missouri. A copy of the transcript of all discussions held during the conference is 

filed in Docket No. 23494, and may be examined by interested persons. 

At that conference a commenter on the recommendation noted that it dealt with 

failures that cause hazards which do not have catastrophic potential, but does not 

deal with hazards which are potentially catastrophic. As noted in the above 

discussion, in the development of requirements for small airplane systems it was 

originally recognized that failures of systems could produce hazards and the earlier 

requirements addressed protection from those failures, but did not address safetycritical systems whose failures could be potentially catastrophic or would be 

catastrophic. As proposed, Section 23.1309 addresses all levels of hazards and the 

proposed requirements are based on the criticality of the system. 

This would be accomplished by requiring all of the airplane's systems to be reviewed 

to determine (1) if the airplane is dependent upon a system function for continued 

safe flight and landing, and (2) if a failure of any system on the airplane, not limited 

to VFR conditions, would significantly reduce the ability of the crew to cope with the 

adverse operating conditions. For airplanes that do not include systems which 

perform either of these safety critical functions, the single-fault or fail-safe concept 

requirements would continue to be applicable as proposed in Section 23.1309(a). 

If the design of the airplane includes systems that perform a function that is needed 

for continued safety flight and landing of the airplane and, accordingly, whose failure 

could be catastrophic, the systems would be required to meet standards that establish 

that failures of the system must be extremely improbable. In addition, on airplanes 

designed for any type of operation other than VFR, the systems whose failures would 

significantly reduce the airplane's capability, of the ability of the crew to cope with 

the adverse operating conditions and, thereby, be potentially catastrophic would be 

required to meet standards that establish that failures of these systems are 

improbable. This standard is applicable if a system failure would reduce the 

capability of the airplane or the crew to cope with adverse operating conditions. It 

was recognized that any failure will reduce the airplane's or crew's capability by 

some degree, but that reduction may not be of the degree that will make operation of 

the airplane potentially catastrophic. The intent of this proposed standard, Section 

23.1309(b), is to have those systems whose failure would be catastrophic or 

potentially catastrophic be evaluated using the latest available techniques and, 

thereby, better assure that failures of such systems will not occur. 

At the conference, a commenter expressed the opinion that the language of 

conference proposal 434 would require all system items to meet the analysis test 

206 
04/12/05 AC 23-17B 

and proof of compliance means. This commenter noted that simple and conventional 

systems can be assessed on the basis of service experience and engineering judgment 

and noted that low performance, simple designed airplanes should be able to use the 

existing method of determining compliance. The FAA agrees with this commenter 

and, as previously stated, proposed Section 23.1309(a) is structured to allow the use 

of existing procedures for simple airplane system designs. 

Another conference commenter expressed a concern over the applicability of Section 

XX.1309 of any of the airworthiness parts. To clarify the applicability of Section 

23.1309, a definition of "system" is included in proposed Section 23.1309(f). 

One commenter noted the difference in the way Part 23 and Part 25 airplanes are 

used and suggested different probability values for Part 23, but did not provide 

recommended values. In reviewing this comment, it should be noted that the 

probability terms do not have values assigned in these proposed requirements of 

other airworthiness parts. Extremely improbable failure conditions have been 

defined in FAA guidance material as those which are so unlikely that they are 

expected to occur within the total operational life of all airplanes of one type. If such 

a definition were to be applied to items of structure, a failure that would cause the 

loss of a wing would be a type of catastrophic failure that would not be expected to 

occur in the life of all airplanes of one type. As cited, current requirements 

applicable to other portions of Part 23 airplanes, such as structures, establish a level 

of safety that does not permit the occurrence of catastrophic failures and, 

accordingly, there is no justification for allowing a lower level of safety for possible 

catastrophic system failures. Accordingly, the proposed language of this proposal 

uses the term "extremely improbable" to define this critical type of failure condition. 

Less critical failure condition terms used in airworthiness requirements for other 

categories of aircraft are also included in this proposal. 

Other review conference comments, not limited to any one commenter, questioned the 

applicability of the rational method for analysis to two-to-four-place airplanes. 

While this proposal allows the continued use of existing certification procedures for 

certification of simple airplane designs, it would, however, also require the use of 

rational procedures if the airplane's design includes systems for the accomplishment 

of safety-critical functions. Because the proposed additional requirements of this 

proposal are added in such a manner as to make their use dependent on the 

complexity of the affected design and because the degree of reliability required in a 

particular design will depend upon the criticality of the system function, there is no 

reason to limit these requirements to a size of airplane or to differentiate between 

single or multiengine types of airplanes. 

One review conference commenter questioned the power supply requirements of the 

conference proposal number 434. This commenter asked if this proposed 

207 
AC 23-17B 04/12/05 

requirement is related purely to electrical power, and if it is, shouldn't it be located 

elsewhere in Part 23. During the discussion of this question, it was pointed out by 

FAA panel members that the proposed requirement does not specify an "electrical" 

power supply and, therefore, would be applicable to any form of power supply 

provided for a system that is required by this chapter. No additional language has 

been added to identify the types of power supply. The proposal language, which is 

similar to that used in other airworthiness parts, identifies requirements for power 

supplies used for each item of equipment, system, and installation required by this 

chapter. By its applicability, it is clear that the requirements are not limited to 

electrical power supplies. 

Additional discussions at the conference suggested that current Section 23.1309 was 

adequate and that the only need was an Advisory Circular (AC) to identify acceptable 

means of compliance. At that time, the FAA made it known that was not the case, and 

cited instances in which a complex safety-critical system had been used in the design 

of small airplanes that could not be properly evaluated under the existing "single 

fault" concept of Section 23.1309. The FAA further noted that the number of 

occurrences where such a complex safety-critical system was being used in the design 

was increasing rapidly and noted the need to revise the requirements to keep them 

current with this rapid expansion of technology being applied to the design of small 

airplanes. 

The FAA attempted to address some current issues relative to electronic flight 

instrument systems and autopilot monitors and limiters in advisory circulars. FAA 

review of resulting material in the draft advisory circulars determined that the 

contents were rulemaking in nature and not suitable in an advisory circular. The 

concern was relative to the complexity and criticality of such equipment. This 

proposal, when adopted, will provide a regulatory basis for determining the 

criticality level of such systems and require corresponding levels of reliability 

requirements which are based on the criticality of the system's function and will 

provide the updated standards needed for the certification of complex safety-critical 

systems in small airplanes. 

This amendment is explained by Final Rule, Docket 25812 as follows: This proposal 

would retain the existing reliability requirements of current Section 23.1309 for 

airplane equipment, systems, and installations that are not complex and do not 

perform safety-critical functions. For those cases where the applicant finds it 

necessary or desirable to include complex, safety-critical systems, this proposal also 

would provide additional requirements for identifying such equipment, systems, and 

installations and would define additional requirements needed for their certification. 

This proposal would permit the approval of more advanced systems having the 

capability to perform critical functions and whose failure condition would prevent the 

continued safe flight and landing of the airplane. 

208 
04/12/05 AC 23-17B 

Two commenters offer comments on proposed Section 23.1309. One of these 

commenters concurs with the concept of updating the reliability requirements 

applicable to airplanes not limited to Visual Flight Rules (VFR) flight, but does not 

concur with this updating for all airplanes. As discussed in Notice Number 89-6, this 

proposal addresses the systems installed on airplanes and is not limited to the 

operations approval of the airplane. The airworthiness standards, as adopted in 

Section 23.1309(a), are based on single-fault or fail-safe concepts and experience 

based on service-proven designs and engineering judgment. These requirements 

should be used for airplanes whose systems are not complex and do not perform 

safety-critical functions. Therefore, Section 23.1309(a) is structured to allow the use 

of existing procedures for simple airplane system designs. 

If the design of the airplane includes equipment, systems, and installations that 

perform functions whose failure condition would prevent continued safe flight and 

landing of the airplane, the occurrence of each failure conditions must be extremely 

improbable. In addition, on airplanes designed for any type of operation not limited 

to VFR, the systems whose failure conditions would significantly reduce the 

airplane's capability, or the ability of the crew, to cope with the adverse operating 

conditions must be improbable. It was recognized that any failure would reduce the 

airplane's or crew's capability by some degree, but that reduction may not be of the 

degree that would make operation of the airplane potentially catastrophic. The intent 

of Section 23.1309(b) is to require that systems whose failure would be catastrophic 

or potentially catastrophic be evaluated using the latest available analysis 

techniques. 

Although future airplane designs limited to VFR operations are not likely to include 

equipment, systems, and installations whose failure condition would prevent 

continued safe flight and landing of the airplane, the applicability of this 

requirement, as discussed above, will provide airworthiness standards if the 

applicant elects to include such systems in the airplane's design. Therefore, the 

applicability of this requirement has not been revised as suggested by this 

commenter. 

One commenter suggests that the critical environmental system considered in Section 

23.1309(c) would be better defined by removing the words "such as" from the 

proposed paragraph and replacing them with the word "including." The FAA agrees 

that the suggested wording more accurately identifies the intent of this paragraph, as 

discussed in this notice. The wording of paragraph (e) of Section 23.1309 has been 

revised accordingly. 

This same commenter notes that there are proposals being considered for a new 

Sections 25.1315 and 15.1317, which deal with the effects of lightning and external 

high energy radiated electromagnetic fields, and suggests that similar actions be 

considered for part 23 rules. Although this comment is beyond the scope of the 

actions proposed in Notice Number 89-6, the FAA recognizes the desirability of 

209 
AC 23-17B 04/12/05 

having the various airworthiness standards address like requirements in the 

respective sections and will consider this comment in future rulemaking actions.” 

Lightning Protection 

See AC 23.1309-1C, “Equipment, Systems, and Installations in Part 23 Airplanes,” 

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”, 

November 1, 1999. 

SAE ARP 5413, “Certification of Aircraft Electrical/Electronic Systems for the 

Indirect Effects of Lightning,” (it is being incorporated into an AC), May 14, 2002. 

SAE ARP 5414, “Aircraft Lightning Zoning”, December 1, 1999. 

SAE ARP 5415, “Users Manual for Certification of Aircraft Electrical/Electronic 

Systems for the Indirect Effects of Lightning”, May 14, 2002. 

SAE ARP 5577, "Aircraft Lightning Direct Effects Certification”, September 30, 

2002. 

High Intensity Radiated Fields (HIRF) 

Special conditions will still be required for critical systems for 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. 

210 
04/12/05 AC 23-17B 

Electronic Engine Control (EEC) Systems 

Section 23.1309 does not apply to powerplant systems provided as part of the 

certificated engines according to paragraph (f). This rule did not envisage digital 

engine controls. Special conditions will be required for electronic engine controls 

without a hydromechanical back-up system that has the reliability and performance of 

a traditional reciprocating engine without electronic engine controls. The special 

conditions will require compliance to HIRF and § 23.1309(a) through (e), according 

to 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 either engine mounted and engine certificated Full Authority Digital 

Electronic Engine Control (FADEC) or EECs 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 either by the engine manufacturer, FADEC, or EEC 

manufacturer should be validated on each software change of 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. 

Standardization of Application of 14 CFR, Part 23, § 23.1309, Regarding 

Hazardous Misleading Heading Information for Attitude-Heading Reference 

Systems (AHRS) 

Purpose 

The purpose of this guidance is to clarify FAA certification policy on the application 

of AC 23.1309-1C, “Equipment, Systems, and Installations in Part 23 Airplanes,” 

regarding hazardous misleading heading information. 

The issue in question is specifically about the application of AC 23.1309-1C for an 

airplane with a certification basis of amendment 23-41 or later. This clarification is 

limited to installations approved for operation in Instrument Meteorological 

Conditions (IMC) under IFR. 

For operations in Visual Meteorological Conditions, a misleading heading indication 

is not considered hazardous. 

211 
AC 23-17B 04/12/05 

Current Regulatory and Advisory Material 

14 CFR part 23, § 23.1309, Equipment, systems, and installations 

Technical Standard Order (TSO) TSO-C151b, “Terrain Awareness and Warning 

System”, May 6, 2002. 

AC 23.1309-1C, “Equipment, Systems, and Installations in Part 23 Airplanes”, 

March 12, 1994.

AC 23-18, “Installation of Terrain Awareness and Warning System (TAWS) 

Approved for Part 23 Airplanes”, June 14, 2000. 

Summary 

A Functional Hazard Assessment (FHA) and the related safety assessments should be 

made for the specific airplane type and configuration. There can be many 

combinations of failures, various mitigating factors, and other functions available to a 

pilot. Because of the many factors that affect the criticality of the heading indication, 

a safety assessment process using AC 23.1309-1C should be used to classify the 

failure conditions for misleading heading information. 

A hazardously misleading heading is usually when the accuracy error is greater than 

10 degrees on the primary heading instrument and it is an undetected error from the 

AHRS. The safety assessment process should consider appropriate mitigating factors 

that might reduce failure conditions effects, including: 

availability, accuracy, and reliability of magnetic compass; 

independent heading instrument; 

independent navigation information; 

moving map if approved for IFR operations; 

TAWS, if approved (for design and performance) according to TSO-C151ab, Class A 

or B, and following AC 23-18 (for airworthiness approval for the installation); 

air traffic control monitoring of the aircraft; and 

exposure time. 

212 
04/12/05 AC 23-17B 

Safety Assessment Process 

A hazardously misleading heading is usually when the accuracy error is greater than 

10 degrees due to a malfunction on the primary heading system using the AHRS. 

The normal accuracy should be within the guidelines under section 23.1301 in this 

AC. For Class I part 23 airplanes, the failure condition for misleading hazardous 

heading is usually considered major according to AC 23.1309-1C. 

For a major failure condition on Class I airplanes, the likelihood for the system to 

provide hazardously misleading heading information to the display should be less 

than or equal to 10-4 per flight hour because of undetected or latent failures. For 

Class II, III, and IV airplanes, the probability for the system to provide hazardously 

misleading heading information to the display should be less than or equal to 10-5 per 

flight hour because of undetected or latent failures. For major failure conditions, a 

qualitative analysis is usually sufficient. 

NOTE 17: There is a difference between hazardous as used in general policy or 

regulations and hazardous failure condition as used in an FHA. When the term

"hazard" or "hazardous" is used in general policy or regulations, it is generally 

defined as follows: 

"Any condition that compromises the overall safety of the airplane or that 

significantly reduces the ability of the flight crew to cope with adverse operating 

conditions." 

When used in an FHA, "hazardous failure condition" is one of the five different 

failure conditions classifications in AC 23.1309-1C. Hazardous failure conditions 

would reduce the capability of the airplane or the ability of the crew to cope with 

adverse operating conditions to the extent there would be the following: 

A large reduction in safety margins or functional capabilities; 

(1) Physical distress or higher workload such that the flight crew cannot be relied on 

to perform their tasks accurately or completely; or 

(2) Serious or fatal injury to an occupant other than the flight crew. 

Mitigating Factors 

During the safety assessment process, mitigating factors should be considered if they 

are independent of the AHRS. These factors should provide additional qualitative 

and quantitative credit and alleviate failure conditions. The mitigating factors may 

include the following: 

availability, accuracy, and reliability of magnetic compass; 

213 
AC 23-17B 04/12/05 

independent heading instrument; 

independent navigation information; 

moving map if approved for IFR operations; 

TAWS, if approved (for design and performance) in accordance with TSO-C151ab, 

Class A or B, and following AC 23-18 (for airworthiness approval of the installation); 

air traffic control monitoring of the aircraft; and 

exposure time. 

Some of these factors may not be appropriate for higher performance and complex 

airplanes. 

Normally, for Class I and some Class II airplanes, the pilot should be scanning the 

instruments for cues to determine misleading information. Therefore, the locations of 

primary instruments and secondary instruments should be within the primary and 

secondary fields-of-view. This is defined in GAMA Publication No. 10, 

“Recommended Practices and Guidelines for Part 23 Cockpit/Flight Deck Design,” 

dated September 2000. 

Magnetic Compass or Independent Heading Instrument 

The usability of the magnetic compass, as a mitigating factor, is dependent on the 

airplane installation on a case-by-case basis. The following guidelines should be 

considered. A magnetic compass or independent heading instrument is considered 

acceptable if the accuracy is within 10 degrees and it is located in the secondary fieldof-view. However, the compass may not be usable if it is susceptible to large errors 

induced by other onboard equipment. For example, an airplane approved for known 

icing may have an unusable compass when a windshield heater is affecting the 

compass; therefore, additional mitigating factors may be necessary to meet the 

appropriate safety level. 

One general concern is that Class II airplanes may have some systems that can affect 

the magnetic compass that are not typically found in Class I airplanes. Therefore, the 

magnetic compass should be usable for many Class I airplanes, but it may not be 

usable for many Class II airplanes or higher. An evaluation should be conducted to 

determine if the magnetic compass is usable and accurate within 10 degrees.

214 
04/12/05 AC 23-17B 

Navigation Information 

Credit should also be given for moving maps displaying position information from

independent navigation systems such as GPS that are approved for IFR operations. 

The moving map should depict the airplane relative to the desired navigation track 

and the navigational tracks heading. 

Navigation systems such as IFR-approved GPS systems can provide an independent 

relative heading source that can be used to verify the primary heading indication. 

Besides, the FAA expects that pilots will naturally crosscheck their moving map 

displays because they are easier to use for navigation than traditional deviation 

needles. 

The moving map should be in the primary field-of-view (preferably) or the secondary 

field-of-view to get credit for the safety assessment process by AC 23.1309-1C. 

TAWS 

TAWS with a display provides the pilot with sufficient information and alerts to 

detect potentially hazardous terrain situations so the pilot may take effective action to 

prevent CFIT events. Although a TAWS with a display is not approved for 

navigation under normal circumstances, under emergency conditions TAWS provides 

acceptable situational awareness to prevent CFIT. 

TAWS may not be usable to determine misleading heading information, but TAWS

should reduce or prevent a CFIT accident, which will mitigate the failure effect. 

TAWS that meet TSO-C151ab, Class A or B, and AC 23-18 should be acceptable 

when pilot confidence proficiency is achieved through service history. 

Air Traffic Control Monitoring 

For most IFR operations, the airplane will be in controlled airspace and under radar 

coverage. Air Traffic Control (ATC) has been the primary mitigating factor for 

heading indicator accuracy for many decades. ATC maneuvers airplanes based on a 

ground track. Heading errors in traditional mechanical Directional Gyros (DG) are 

still identified by ATC when pilots do not maneuver in the direction ATC expects. 

Most of the time ATC will identify heading errors when the pilot fails to adequately 

update their DG or their horizontal situation indicator has failed. 

However, there are locations at some airports where radar coverage is not available. 

There is usually some type of navigation signal that pilots can use to get into and out 

of the airport. These are the types of departures and approaches where heading 

accuracy may be important. Although ATC may be a mitigating factor, specific 

guidelines are not being provided now. 

215 
AC 23-17B 04/12/05 

Exposure Time and Risk Assessment 

Exposure time and risk factors that may be considered are as follows: 

automatic pre-engagement self test; 

monitoring capabilities; 

maintenance intervals; and 

phase of flight, and so forth. 

In the SAE Standard ARP4761, “Guidelines and Methods for Conducting the Safety 

Assessment Process on Civil Airborne Systems and Equipment,” exposure time is 

defined as follows: 

"The period of time between when an item was last known to be operating properly 

and when it will be known to be operating properly again." 

In this guidance, risk is also being considered. 

Risk assessment is the composite of: 

the outcome or effect of the failure condition; 

the average probability of the failure condition per flight hour; and 

exposure time; and 

likelihood of the outcome/effect of the hazard will occur. 

the exposure time. 

For risk assessments, a credit should be factored into the analysis that is proportional 

to the probability that a given failure condition will occur on the aircraft. This credit 

is related to the system level of the failure condition, the operational and 

environmental conditions, and the phase of flight. 

To determine exposure time, the time the pilot would only be using heading

information to navigate should be considered. Only a percentage of a typical IFR 

flight is flown navigating solely on the heading instrument. The critical times for 

heading are when the pilot is not navigating with a navigation instrument but is only 

using the heading instrument. 

216 
04/12/05 AC 23-17B 

The details on how to calculate the “Average Probability Per Flight Hour” for a 

Failure Condition are given in Appendix C of AC 23.1309-1C. A common 

assumption for 14 CFR, part 23 airplanes is the average flight duration is one hour. 

The FAAs General Aviation and Air Taxi Activity Survey gives estimated flight 

hours for IMC operations for specific airplanes. When researching the operations of 

specific airplane models, the result indicates that Class I airplanes are operated in 

IMC conditions about 20 percent of their total flight hours and Class II airplanes 

about 30 percent. Also, the FAA believes that only portions of most IFR flights are 

flown by navigating solely on the heading instruments. 

Using the percentages for “phases of flight” from a Boeing accident study, the FAA 

estimates that on any given IFR flight, navigation solely by the heading instrument 

would only occur about 20 percent of the time. For this analysis, Class III and Class 

IV airplanes are assumed to be operating IFR nearly 100 percent of the time.

Acceptable Configurations for IFR Use 

Some acceptable configurations for IFR approval include: 

Configuration A: 

One dependable magnetic compass. 

(1) A reliable magnetic gyroscopically stabilized heading system that together 

meets the AC 23.1309-1C reliability levels, such as 10-4 per flight hour for Class I 

airplane. 

Configuration B: 

A magnetic compass that meets the airworthiness requirements but cannot be 

given credit for AC 23.1309-1C purposes for the reasons under Magnetic 

Compass or Independent Heading Instrument paragraphs in the Safety 

Assessment Process. 

(2) One magnetic gyroscopically stabilized heading system that meets the AC 

23.1309-1C reliability levels, such as 10-4 per flight hour for Class I airplane, alone or 

in combination with one or more of the mitigating factors addressed such as moving 

map or TAWS with a display. 

Configuration C: 

(1) A magnetic compass that meets the airworthiness requirements but cannot be 

given credit for AC 23.1309-1C purposes for the reasons stated under Magnetic 

Compass or Independent Heading Instrument paragraphs in the Safety Assessment 

Process. 

217 
AC 23-17B 04/12/05 

(2) Two magnetic gyroscopically stabilized heading systems that together meet 

the AC 23.1309-1C reliability levels, such as 10-4 per flight hour for a Class I 

airplane. 

Configuration D: 

A magnetic compass that meets the airworthiness requirements but cannot be 

given credit for AC 23.1309-1C purposes for the reasons stated under Magnetic 

Compass or Independent Heading Instrument paragraphs in the Safety Assessment 

Process. 

(1) Two magnetic gyroscopically stabilized heading systems with each being less 

reliable than required by AC 23.1309-1C (such as 10-4 or 10-5). 

(2) A comparator monitor. In this case, the warning from the comparator monitor 

would alleviate a hazardous misleading heading. However, if the system still has a 

reliability problem, many nuisance alarms would cause the pilot to ignore the 

warnings. 

Conclusion 

There are many factors that affect the criticality/severity of hazardously misleading 

heading indication, so a safety assessment process using AC 23.1309-1C should be 

used to classify this failure condition. 

The safety assessment process should consider appropriate mitigating factors that 

might alleviate failure conditions, including: 

availability and reliability of magnetic compass; 

independent heading instrument; 

navigation information; 

moving map; 

terrain awareness warning system; 

air traffic control monitoring of the aircraft; and 

exposure time. 

218 
04/12/05 AC 23-17B 

Amendment 23-49 

A proposed revision to NPRM 94-21 explains this amendment as follows: 

“Proposed new Sec. 23.1309(a)(4) would correct an inadvertent omission that 

occurred when the FAA issued Amendment No. 23-41 (55 FR 43306, October 26, 

1990). The omitted requirement was adopted by Amendment No. 23-34 as a 

portion of Sec. 23.1309(d) and read: "In addition, for commuter category 

airplanes, system and installations must be designed to safeguard against hazards 

to the airplane in the event of their malfunction or failure." (52 FR 1833, January 

15, 1987.) To correct this oversight, and to continue the single fault provision of 

this paragraph, Sec. 23.1309(a)(4) is being proposed.” 

219 
AC 23-17B 04/12/05 

INSTRUMENTS: INSTALLATION 

23.1311 Electronic display instrument systems 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.13.

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, 

except that independent secondary attitude and rate-of-turn instruments that comply 

with § 23.1321(a) if the primary electronic display instrument system for a pilot 

presents this information are required. Therefore, both electronic attitude and 

heading instruments can use the same attitude and heading reference source. An 

applicant for a single AHRS can present their rationale, including a hazard analysis, 

explaining how their single AHRS system can meet reliability requirements for a 

possible catastrophic failure. For a single AHRS, the AFM 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

A proposed revision to NPRM 89-6 has the following explanation: “A significant 

number of electronic display systems have become available for installation in small 

airplanes. These systems include display of all parameters that are typically 

displayed on a small airplane instrument panel. Approval of these systems in small 

airplanes was addressed by several conference proposals that proposed to amend 

Sections 23.1303, 23.1305, 23.1321, 23.1323, and 23.1337. 

Conference proposal 420 from the General Aviation Manufacturers' Association 

(GAMA) recommended removing the words "instrument" and "indicator" from 

Section 23.1303 and require specific "data" be displayed rather than require specific 

"instruments". Conference proposal 428 recommended amending Section 23.1305 by 

changing "powerplant instruments" to "powerplant displays" and "indicators" and 

"indicating" to displays". Conference proposal 436 recommended amending Section 

23.1321 by changing "instrument" to "display" or "data on the flight displays". 

Conference proposal 439 recommended amending Section 23.1323 by changing 

"instrument" and "indicator instrument" to "display". Conference proposal 450 

220 
04/12/05 AC 23-17B 

recommended amending Section 23.1337 by adding a new paragraph (a) to state 

"Displays other than individual indicators may be used if it is shown that adequate 

isolation is provided between engines and engine parameters." 

The basic justification given for all of the recommended changes is that the current 

requirements of Part 23 were written when the required data could only be supplied 

using individual instruments. New technology now allows this same data to be 

displayed in a different manner, possibly with all the data on a common display. 

It is desirable to take advantage of available new technology. The benefits include 

safer (less prone to misreading) displays with less cockpit space and of equal or 

lower cost than the cluster of typical individual instruments. As technology advances, 

the amount of energy used for these displays and the cost can be further reduced. 

The FAA agrees that when current requirements of Part 23 were written, only 

mechanical or electromechanical instruments that functioned independently for each 

parameter displayed were envisioned. The engine instruments and systems 

envisioned were isolated and independent. A single failure of any engine or any of its 

systems could not affect the operation of any other engine. The requirements were 

based on "single fault" or "fail safe" concepts and, when these current requirements 

were promulgated, the FAA did not envision use of complex, safety-critical systems in 

small airplanes. All envisioned instruments were single function; i.e., a failure would 

cause loss of only one instrument function, although several instrument functions may 

have been housed in a common indicator case. 

Since the conference, the FAA has further studied the problems associated with 

installation of current technology indicators in small airplanes. As is discussed 

relative to amending Section 23.1309, these current technology systems have 

potential for being critical for continued safe flight of the airplane. The potential for 

increased clarity in data display and the concentration of data displays in a single 

indicator increases the potential criticality of failures. It is anticipated that pilots 

using these new instrument systems will become increasingly dependent on the use of 

them because of the tasks they perform for the pilot. After a period of time, where 

these electronic indicators are located in the primary instrument panel locations, it is 

anticipated that pilots will find it more difficult to transition to back-up or secondary 

indicators when failure occurs, such as reverting to use of needle-ball and airspeed 

for airplane attitude control when the artificial horizon instrument system fails. 

The electronic indicators are expected to have significantly different modes of failure 

where they go from performing perfectly to total failure, whereas the mechanical and 

electromechanical indicators typically deteriorated in performance over a period of 

time such that they were replaced before a total failure that prevented them from 

providing useful information to the pilot. 

221 
AC 23-17B 04/12/05 

Current technology instrument systems with electronic indicators for small airplanes 

may very considerably in functional capability, complexity, and cost. Due to the 

economic considerations, the most expensive, complex, and reliable electronic 

instrument systems will only be installed in airplanes fitting a like description. 

The electronic instrument systems can readily provide digital indication of exact 

numbers, moving pointer on a scale, and various other formats and combinations of 

them all. The FAA is especially concerned that pilots be provided adequate sensory 

cues as to whether numbers displayed are increasing or decreasing and how fast they 

are changing. Also of concern is that digital indication may not show the normal 

operating range cues to direction or rate of change or operational limits. 

As a result of these concerns and this further review of possible ways to address 

electronic instrument systems in Part 23, the FAA concluded that a new Section 

23.1311 for these systems is better than amending several sections, which may result 

in an unclear treatment of the issue. 

Relative to identifying the indicating means of the electronic instrument system, the 

FAA has reviewed existing materials and functions to be performed and has 

concluded the proper identifier is "indicator" rather than "display" as recommended. 

Sections 23.1303(a), (b), and (c) require basic flight and navigational instruments for 

small airplane certification. These mechanical instruments have performed their 

intended function very well over the years and these basic instruments will remain 

necessary for safety even when the current technology systems are installed. 

Therefore, this proposal will allow displacement of these instruments from the 

primary location for such instruments for compliance with the requirements of 

23.1321, provided their location in a secondary location is such that they are usable 

and in compliance with Section 23.1321(a) requirements. It is the FAA's intent that 

this will continue the requirements that airspeed, altitude, and magnetic compass 

information will remain available to the pilot after total failure of the airplane's 

electrical power system. 

This proposal will allow electronic display indicators for engine parameters without 

isolation and independence of engine instruments as is now required. In developing 

this proposal the FAA considered the operational characteristics of airplane engines, 

the proposed amendment to Section 23.1309 in this notice for assessing failures and 

their consequences, and the cues available to the pilots for assuring an engine 

instrumentation failure would not create a condition where the pilot would encounter 

significant difficult in operating the engines. 

Due to the dependence pilots are expected to place on use of electronic indicators 

when the indicators include information essential to airplane attitude control, the 

FAA is proposing secondary attitude and rate-of-turn instrument systems that comply 

222 
04/12/05 AC 23-17B 

with Section 23.1321(a) when the electronic indicators include display of attitude and 

rate-of-turn. 

Electronic indicator legibility is expected to change as the cathode ray tubes (CRT)

used in the electronic indicators age. Therefore, it is considered necessary that 

instructions for continued airworthiness relative to the useful life be addressed in 

compliance with Section 23.1529. 

Electronic indicator systems will have great potential for inhibiting information to 

maximize the effect of other information in various phases of flight. Attitude, 

airspeed, altitude, and powerplant parameters needed to set power within established 

limits are information the FAA has concluded must be displayed during all normal 

modes of operations and, therefore, may not be inhibited during normal modes of 

operation. Information that is considered essential to continued safe flight must 

remain available on indicators usable by the pilot after any single failure or 

combination or probable failures without need for immediate crew action. At a 

minimum, without considering specific characteristics of an airplane's design, 

attitude, airspeed, and altitude must remain available without any crew action after 

such a failure, whereas a failure that would remove other essential information from 

displays, without resulting in an immediate hazard, would be acceptable provided the 

essential information could be returned to a usable indicator in a safe elapsed time.” 

This amendment is explained by Final Rule, Docket 25812, as follows: “One 

commenter asks if the wording of proposed Section 23.1311(c), concerning electronic 

display indicators with features that make isolation and independence between 

powerplant instrument systems impractical, will be supported by an appropriate 

amendment to require such isolation. As discussed in Notice No. 89-6, the current 

requirements of part 23 address powerplant instruments that could provide the 

required data only by using individual instruments. Accordingly, the isolation and 

independence referred to in Section 23.1311(c) are currently required in Section 

23.903(c). The objective of this regulation is to allow the use of electronic display 

indicators that will not provide the isolation and independence considered in the 

current requirements. The FAA is not considering an additional amendment to 

address this issue.” 

See AC 23.1311-1A, “Installation of Electronic Displays in Part 23 Airplanes,” for 

further guidance. 

Altitude 

Digital-Only (alphanumeric) displays for barometric altitude should not be approved. 

223 
AC 23-17B 04/12/05 

Amendment 23-49 

A proposed revision to NPRM 94-21 explains this amendment as follows: “This 

proposal would revise Sec. 23.1311 to remove redundant requirements and to clarify 

which secondary instruments are required and the visibility requirements for these 

instruments are required and the visibility requirements for these instruments. When 

Sec. 23.1311 was adopted by Amendment No. 23-41 (55 FR 43306, October 26, 

1990), several nonsubstantive changes were made to the proposals in Notice No. 89-6 

(54 FR 9345, March 6, 1989) to remove the redundancy included in the notice. In the 

process certain provisions, such as the one that permitted the installation of 

mechanical secondary instruments, were inadvertently omitted from the final rule. 

Since the final rule, discussions with airplane manufacturer representatives have 

shown that the requirements defining the instrument panel location where secondary 

instruments may be installed are also not clear. Accordingly, the FAA is proposing 

to revise this section to correct and clarify these portions. 

Current Sec. 23.1311(a), which requires electronic display indicator installations 

that are independent to each pilot station, would be deleted because it is redundant 

with Sec. 23.1321(a). Section 23.1321(a) requires that each flight, navigation, and 

powerplant instrument for use by any required pilot shall be located so that any pilot 

seated at the controls can monitor the instruments with minimum head and eye 

movement. As stated in the preamble of Notice No. 89-6 (54 FR 9345, March 6, 

1989) regarding the proposed revision to sec. 23.1321, "This revision also clarifies 

the rule relative to instrumentation that must be provided for each pilot required for 

type certification or by the applicable operating rules. If a pilot is required by any 

applicable requirement, then that pilot must be provided all instrumentation required 

for any operations for which the airplane is approved." Accordingly, the 

requirements of current Sec. 23.1311(a) would be removed. 

In place of current paragraph (a), proposed Sec. 23.1311(a) would be a revision of 

current paragraph (c) that would clarify what instruments are required and the 

visibility of those instruments. Proposed new Sec. 23.1311(a)(1) would require 

electronic display instrument installations to meet the arrangement and visibility 

requirements of Sec. 23.1321(a). 

Proposed Sec. 23.1311(a) (2), (3), and (4) would be redesignated with no changes 

from current Sec. 23.1311(c)(1), (2), and (3).

Proposed Sec. 23.1311(a)(5) would continue the requirement of Sec. 23.1303(c) for a 

magnetic direction indicator and, in addition, would require either an independent 

secondary mechanical altimeter, airspeed indicator, and attitude indicator or 

individual electronic display indicators for the altimeter, airspeed, and attitude that 

are independent from the airplane's primary electrical power. 

224 
04/12/05 AC 23-17B 

These secondary instruments may be installed in panel positions other than the 

primary location as long as the selected location allows the pilot to properly monitor 

the instruments and control the airplane. 

The substance of proposed paragraph (a)(5) is both a substantive change and a 

combination of the current Sec. 23.1311(b), which states that certain electronic 

display indicators must be independent of the airplane's electrical power system, and 

current Sec. 23.1311(c)(4) which requires independent secondary attitude and rateof-turn instruments and specifies the location of those instruments and specifies the 

location of those instruments. Proposed Sec. 23.1311(a)(5) would delete the 

requirement for a rate-of-turn instrument (in current sec. 23.1311(c)(4)) and specify 

that the required secondary instruments are those that provide altitude, airspeed, 

magnetic direction, and attitude. The information that would be provided by a 

secondary rate-of-turn instrument would not appreciably add to the safe operations 

of the airplane if the pilot has the information provided by the secondary attitude 

instrument. 

Current Sec. 23.1311(b) requires that electronic display indicators required by Sec. 

23.1303(a), (b), and (c) be independent of the airplane's electrical power system. The 

original intent of the requirement for secondary instruments, as stated in Notice No. 

89-6, was to require the installation of either mechanical instruments or independent 

electronic display indicators powered by a source independent of the airplane's 

electrical system. However, the current rule does not clearly state this and does not 

address the installation of mechanical instruments. Proposed Sec. 23.1311(a)(5),

would allow either secondary electronic display indicators or mechanical 

instruments to provide a crew with information essential for continued flight and 

landing in the event of failure in the airplane's electrical power system. 

Current Sec. 23.1311(c) (5) and (6) would be redesignated as Sec. 23.1311(a)(6) and 

(7) without change. 

Proposed new Sec. 23.1311 (b) and (c) would continue the requirements of current 

Sec. 23.1311 (d) and (e) without change. 

This proposal would revise Sec. 23.1311 to remove redundant requirements and to 

clarify which secondary instruments are required and the visibility requirements for 

these instruments. No comments were received on the proposal, and it is adopted as 

proposed.” 

This amendment is explained by Final Rule, Docket 27806 as follows: “This 

proposal would revise Sec. 23.1311 to remove redundant requirements and to clarify 

which secondary instruments are required and the visibility requirements for these 

instruments. No comments were received on the proposal, and it is adopted as 

proposed.”

225 
AC 23-17B 04/12/05 

23.1321 Arrangement and visibility 

The corresponding rules in CAR 3 are CAR 3.661 and 3.662. 

The corresponding rules in the Airship Design Criteria, FAA-P-8110-2, Change 2, are 

sections 6.4 and 6.7.

Original Issue and Subsequent

When applying this rule to powerplant instruments in multiengine airplanes, assure 

there is no confusion about 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

A proposed revision to NPRM 71-13 explains this amendment as follows: “This 

proposal is considered necessary because of the increasing complexity of the new 

high performance airplanes and the increasing volume of air traffic. In addition, the 

proposed new paragraph (d) is in furtherance of the principle of cockpit 

standardization.” 

The NPRM for magnetic direction indicators was revised by Final Rule, Docket 

11011 as follows: “Proposed Sec. 23.1321(d)(4) has been revised to provide that the 

magnetic direction indicator required by Sec. 23.1303(c) need not be mounted in the 

instrument panel position proposed. The proposed requirement might create 

difficulties in calibrating and adjusting those indicators in the proposed position.” 

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 according to §§ 23.1525 and 23.1583(h). 

226 
04/12/05 AC 23-17B 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this amendment as follows: “The 

present visibility requirement in Sec. 23.1321(a) pertains only to flight, navigation, 

and powerplant instruments. Since other devices may be intended for use in flight 

and a pilot may develop vertigo by his head movements to see these devices, the 

location of all visual indicators for use in flight needs to be considered. Of course, 

flight safety will remain the primary consideration in position priority. The standard 

of "minimum practicable deviation" will remain, if the rule is adopted as proposed, 

thus allowing considerations of all factors related to utilization and flight safety. The 

same proposal is also made for Secs. 25.1321(a), 27.1321(a), and 29.1321(a). 

The proposed Sec. 23.1321(e), if adopted, would extend the requirement of the 

present and proposed Sec. 23.1321(a) to require consideration of all possible cockpit 

lighting conditions for visibility of required malfunction indicators. 

The FAA considered the question whether the proposal should apply only to required 

malfunction indicators or to all indicators, and believes that if a malfunction 

indicator has been provided for the crew it should be effective under all cockpit 

lighting conditions.” 

A revision to NPRM by Final Rule Docket, 14625, was in response to comments as 

follows: “Several commentators questioned proposed Secs. 23.1321(a), 25.1321(a), 

27.1321(a) and 29.1321(a) (Proposals 5-3, 5-23, 5-42, and 5-56, respectively)

concerning the visibility of instruments with a visual indicator for use in flight by any 

pilot. It was stated that the proposed standard did not adequately provide for 

instruments that were monitored infrequently during flight as contrasted with 

instruments more critical to flight safety. With respect to rotorcraft, one 

commentator also objected on the ground that the space available for mounting 

instruments was extremely limited. The FAA agrees that priority of instruments 

location needs to be more definitive. Therefore, proposed Secs. 23.1321(a), 

25.1321(a), 27.1321(a) and 29.1321(a) are withdrawn. 

Several commentators objected to the phrase "under all cockpit lighting conditions 

"in proposed Secs. 23.1321(e), 25.1321(e), 27.1321(d), and 29.1321(g), contending 

that it includes everything from lightning to total darkness resulting from complete 

electrical failure at night. In light of these comments and after further FAA review, 

proposed Secs. 23.1321(e), 25.1321(e), 27.1321(d), and 29.1321(g) are revised to 

require consideration of all probable cockpit lighting conditions with respect to the 

visibility of malfunction indicators.” 

227 
AC 23-17B 04/12/05 

Amendment 23-41 and Subsequent 

A proposed revision to NPRM 89-6 explains this amendment as follows: “This 

proposal would require those instruments used during certain maneuvers to be 

located such that minimum eye or head movement is needed to monitor the airplanes 

flight path and these instruments. Powerplant instruments for which the location 

requirements apply would be limited to those needed to set power within powerplant 

limitations. The proposed revision of paragraph (d) would extend the T-arrangement 

of flight instruments to include all small airplanes certificated for all flight under 

instrument flight rules. This revision also clarifies the rule relative to 

instrumentation that must be provided for each pilot required for type certification or 

by the applicable operating rules. If a pilot is required by any applicable 

requirement, then that pilot must be provided all instrumentation required for any 

operations for which the airplane is approved. 

Airplanes certificated to Part 23, in most cases, are certificated with only one 

required pilot. However, many of such airplanes subsequently enter Part 135 

operations where two pilots are required. A significant function of the second 

required pilot is to monitor the airplanes flight path regardless of whether or not this 

second pilot is controlling the flight path. Therefore, this second pilot must have 

flight instruments available that are installed in accordance with all of the criteria of 

Part 23 and the affected operating rules, with specific emphasis on the requirements 

of Section 23.1321.” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this amendment as follows: “The 

proposed revision to Sec. 23.1321(d) would remove the wording that limits the 

instrument location requirement to airplanes certificated for flight under instrument 

flight rules or airplanes weighing more than 6,000 pounds. Instruments are for the 

pilot and should be located near that pilots vertical plane of vision without regard to 

what flight rules are approved for the airplane's operation or the maximum weight of 

the airplane.” 

228 
04/12/05 AC 23-17B 

23.1322 Warning, caution, and advisory lights 

Policy is available in AC 23.1311-1A. 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.8. 

Amendment 23-17 and Subsequent 

A proposed revision to NPRM 75-10 added this rule with an explanation as follows: 

“The use of lights as sources of information in airplanes is common practice. The 

FAA considers that standardization of the colors of these lights is an appropriate 

extension of cockpit standardization. Parts 27 and 29 currently contain color 

standards and it is proposed that further changes in those standards be made. One of 

these changes, of a clarifying nature, that should be noted is in the description of 

warning lights as contained in proposed paragraph (a). This proposal is one of four 

identical proposals affecting Secs. 23.1322, 25.1322, 27.1322, and 29.1322. These 

proposals provide standardized light requirements and also provide for approved 

variance from the colors specified for lights such as marker beacon lights, if found to 

be appropriate. It should be noted that the proposals also specify blue as the color 

for position indication, agreement, and correct response lights, blue being a color 

currently used successfully in service.” 

Amendment 23-17 was revised by Final Rule, Docket 14324, to add the following: 

“in the cockpit” in the heading and removed the proposed requirement for blue 

lights. The explanations in the final rule are: “One commentator suggested a 

clarification of the lead-in of the proposal to limit its applicability to lights installed 

in the cockpit as indicated in the explanation to the proposal. The FAA agrees, and 

the lead-ins of Secs. 23.1322, 25.1322, 27.1322, and 29.1322 have been clarified” 

and “The commentator noted that requiring a blue light for position indication was 

not always appropriate since blue was difficult to see in direct sunlight but was 

readily distinguishable in heavily shaded installations. The FAA agrees that blue 

should not be an established standard applicable to all installations. Therefore 

proposed Secs. 23.1322(d), 25.1322(d), 27.1322(d) and 29.1322(d), concerning blue 

lights, are withdrawn.” 

229 
AC 23-17B 04/12/05 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal is addressed in conference proposal 438. It requires that the specific colors 

be consistent with change in brightness over the full range of ambient light conditions 

in the cockpit and that the luminance difference and/or color difference be sufficient 

to preclude confusion or ambiguity under all probable cockpit lighting conditions. 

Light color is not controlled by a lens color cathode ray tube (CRT) displays now 

being incorporated into airplanes. Cockpit lighting evaluations are required in 

Section 23.1321 and clarification if needed in this section to assure compliance with 

these requirements.” 

This amendment clarifies Final Rule, Docket 26344 as follows: “This proposes to 

amend Sec. 23.1322 to require the warning, caution, and advisory lights to be 

effective under all probable cockpit lighting conditions. One commenter 

recommended the words "all probable cockpit lighting conditions" be revised to "all 

normal cockpit illumination." The reason given for the recommendation is that the 

proposed words could include the need to consider a blinding lightning flash. 

The FAA disagrees. If conditions exist where a "blinding" lightning flash occurs, 

none of the lights will be visible while the pilot(s) are blinded and this would not be 

considered to be a probable light condition. The lights should be evaluated for the 

lighting conditions that will occur immediately after that flash to ensue that, as 

quickly as vision is restored after the exposure to the blinding flash, they will provide 

effective warnings, cautions, and advisories. As noted in the NPRM, these lights need 

to be consistent over a full range of ambient light conditions. The words used in this 

proposal describe the need to evaluate the lights over this range of light conditions. 

One commenter supports the proposal. This proposal is adopted as proposed.” 

See AC 23.1311-1A for additional guidance for electronic displays. 

230 
04/12/05 AC 23-17B 

23.1323 Airspeed indicating system 

The corresponding rule in CAR 3 is CAR 3.663. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.9.

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 four-pound bird strike, the minimum distance between pitot tubes that can be 

accepted is 14 inches, measured in a straight line. 

Amendment 23-7 and Subsequent 

A proposed revision to NPRM 67-14 changed five MPH to five knots with an 

explanation as follows: “Nautical units of speed and distance. Consistent with the 

gradual transition from statute miles to nautical miles being made throughout the 

Federal Aviation Regulations, it is proposed, in addition to certain specific changes 

made in this notice, to amend Part 23 by changing all references to "miles" and 

"miles per hour," to "nautical miles" and "knots," respectively, wherever the former 

are used.” 

Amendment 23-17 is additionally explained by Final Rule, Docket 8083, as follows: 

“The amendments are therefore adopted as proposed, using the following criteria: 

(1) Conversions from miles per hour to knots are rounded off to whole units to avoid 

fractions of a knot unless a more accurate measurement is necessary under the 

present rule, and (2) current requirements expressing extremely low airspeeds such 

as 5 or 10 miles per hour, are not changed numerically (such as to 4 or 9 knots, 

respectively) since the substantive difference in these cases is approximately 1 mile 

per hour and is not practically significant.” 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this change to the amendment as 

follows: “Proposed Sec. 23.1545(a) would require airspeed limitation marks be 

located at the corresponding indicated airspeeds instead of at the calibrated 

airspeeds, thereby providing a more useful indicator dial for operation of the 

airplane. However, to ensure uniformity, it is also proposed to revise Sec. 23.1323 to 

require that the airspeed instrument be calibrated according to the standard 

relationship between pressure and airspeed. (See the proposal for Sec. 23.1323.) 

The allowable airspeed system error would then include position error, but exclude 

the airspeed instrument calibration error, consistent with Part 25. The airspeed 

231 
AC 23-17B 04/12/05 

range for calibration, as set forth in proposed Sec. 23.1323, would encompass the 

airspeed limitations in proposed Sec. 23.1545(a).” 

Indicated airspeed was changed to calibrated airspeed in Amendment 23-20 by Final 

Rule, Docket 14625, based on the following: “Another commentator (who concurred 

with the intent of the proposal to revise Section 25.1323) questioned whether the 

allowable error, in proposed Section 23.1323(b), should be expressed in terms of 

calibrated airspeed rather than indicated airspeed and whether the airspeed ranges 

specified in proposed Section 23.1323(b)(1) and (b)(2) were unnecessarily severe 

accuracy requirements. The FAA agrees that the error allowed by proposed Section 

23.1323(b) should be expressed in terms of calibrated airspeed to be in accord with 

normal practices. However, the FAA believes that the airspeed ranges in proposed 

Secs. 23.1323(b)(1) and (b)(2) are needed to encompass the airspeed limitations in 

proposed Section 23.1545. The FAA also believes that, as a practical matter, the 

accuracy requirements in proposed Section 23.1323 are not significantly more 

restrictive than the current requirements.” 

Amendment 23-34 and Subsequent

See AC 23-8B, “Flight Test Guide for Certification of Part 23 Airplanes,” for 

additional guidance. 

A proposed revision to NPRM 83-17 added commuter category rules based on the 

following: “Part 135, App. A., Sec. 13(b) and (d).”

Amendment 23-42 and Subsequent

This amendment added the requirement for a heated pitot tube when a plane is 

approved for flight in instrument meteorological conditions or flight into known icing 

conditions. The NPRM states the following: “Maintaining a functional and 

accurate airspeed system is essential to safe and reliable control of an airplane in 

instrument meteorological conditions and flight in icing conditions.” The proposal 

was considered necessary as a minimum airworthiness standard when operations in 

the airplane are to be approved. Therefore, the use of a drain trap/sump in meeting 

the requirement for a positive means of drainage is allowable, but an applicant must 

consider a failure condition of an overflowing trap/sump that causes misleading 

readings of airspeed. Waiting for an airspeed error to occur from excess moisture is 

not acceptable for an airplane approved for flight in instrument meteorological 

conditions or flight in icing conditions. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this change to this amendment as 

follows: “The proposed new Sec. 23.1323(c) would add a requirement that each 

232 
04/12/05 AC 23-17B 

airspeed indicating system design and installation should provide positive drainage 

of moisture from the system. This proposal is consistent with the provisions required 

for a static system by Sec. 23.1325(b). 

If moisture enters, or accumulates in, an airspeed indicating system, that moisture 

could cause erroneous airspeed indications or the complete loss of airspeed 

information. The resulting loss of accurate airspeed information would be hazardous 

to the operation of the airplane; therefore, to assure the safety of the airplane, the 

FAA would need to apply the more general airworthiness requirements of Secs. 

23.1301 and 23.1309 to such a system and require provisions for drainage of 

moisture. Accordingly, this proposed revision of the airspeed indicating systems 

requirements only clarifies the criteria that must be applied to airspeed indicating 

systems. 

To better organize the requirements that are applicable to the airspeed systems on all 

airplane categories and those that would be additional requirements for the airspeed 

systems of commuter category airplanes, the FAA proposes to redesignate existing 

paragraphs (c) and (e), respectively, as paragraphs (e) and (d). By this 

redesignation, paragraphs (a), (b), (c), and (d) would apply to all airplanes, and 

paragraphs (e) and (f) would include additional requirements applicable to 

commuter category airplanes. 

The proposal for redesignated paragraph (e) would also remove the words "in flight 

and" from the first sentence of that paragraph. This would remove the requirement 

for the airspeed indicating system to be calibrated in flight. The requirement for an 

in flight calibration is provided in paragraph (b). Proposed redesignated paragraph 

(e) would apply to the calibration needed to determine the system error during the 

accelerate-takeoff ground run. 

As identified in the background section of this notice, the FAA is issuing additional 

notices that address proposed changes to the requirement for powerplant, flight, and 

airframe. Proposed revisions to subpart G in the flight NPRM include placing all of 

the requirements for what must appear in the AFM in that subpart. With the 

proposals to revise the AFM requirements, the flight NPRM also proposes that the 

requirement in existing Sec. 23.1323(d) (to show the relationship between IAS and 

CAS in the AFM) be added to Sec. 23.1587 as proposed new paragraph (d)(10). 

Because the AFM requirement would be added Sec. 23.1587, it no longer needs to 

appear in Sec. 23.1323. Accordingly, this notice proposes to remove the text of 

existing Sec. 23.1323(d). 

Proposed new Sec. 23.1323(f) would provide that, on those commuter airplanes 

where duplicate airspeed indicators are required, the airspeed pitot tubes must be 

located far enough apart so that both tubes will not be damaged by a single bird 

strike.” 

233 
AC 23-17B 04/12/05 

23.1325 Static pressure system 

The corresponding rule in CAR 3 is CAR 3.665. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.10.

Amendment 23-1 and Subsequent

This amendment is explained by Final Rule, Docket 1186, as follows: “The notice of 

proposed rule making provided that the influence of airplane characteristics not 

seriously affect the accuracy of instruments having static pressure connections, 

further specified certain static system design and installation details, required a 

system proof test, and made provision for countering icing conditions. 

Section 23.1325(a) has been amended to specify that static pressure connections be 

vented so that external forces least affect instrument accuracy. This change relaxes 

the requirement of the notice and parallels the wording in the comparable section of 

Part 25. 

Two commentators objected to proof testing the static pressure system in each 

production airplane. The rule, however, does not require test of each production 

airplane although that would be one way to demonstrate conformity to the approved 

type design. As suggested by another commentator, sample testing and production 

flight check could be another acceptable means of showing conformity if there were a 

showing of an adequate quality control system and compliance with the equipment 

function and installation requirements of Section 23.1301. 

The Agency does not concur with various comments recommending that it adopt the 

Air Force static system leakage rate at an arbitrary specified altitude with associated 

instruments disconnected. Since many static system leaks occur at instrument 

disconnect points, it is considered that a realistic test for Part 23 airplanes should be 

made with the instruments connected. Section 23.1325(b)(2) has been amended, 

however, to allow a reasonable leakage loss tolerance with associated instruments 

connected at the airplane maximum operating altitude. 

With regard to negating the effect of icing conditions, one commentator 

recommended deleting the requirement altogether in low-performance airplanes 

operating under VFR conditions. Two other commentators recommended that a 

protected alternate source of static pressure be allowed, while a fourth recommended 

use of an optional ice-free static pressure source of less accuracy where calibration 

is given the pilot. 

Since static vent icing may occur during both VFR and IFR conditions with 

hazardous consequences, the Agency believes there is ample justification for antiicing as a certification requirement on all airplanes employing a static pressure 

234 
04/12/05 AC 23-17B 

system for required instruments. In response to the comments, the proposed rule has 

been expanded to permit the use of an alternate static source having a prescribed 

accuracy tolerance, and, where needed, a correction card.” 

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 results. The rule provides for either an anti-icing means or 

an alternate static source as follows: 

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 AFMs 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, except it may have reduced accuracy with compensation from a 

correction card. 

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 23-6 and Subsequent 

A proposed revision to the NPRM 66-44 explains this rule as follows: “One of the 

major objections to the current requirements is the 100-foot per minute leak rate 

specified in Secs. 23.1325(b)(2) and 25.1325(c)(2). In this connection, 

representatives of industry recognized that a 100-foot per minute leak rate at 

maximum cabin differential pressure would result in altitude errors of from 0 to 3 feet 

on most aircraft. Furthermore, on many aircraft, the altitude error resulting from 

leak rates as high as 3,000 feet per minute at maximum cabin differential pressure 

would not exceed 20 feet. It was recommended that a leak rate equivalent to 2 

percent of the maximum cabin differential pressure at the maximum cabin differential 

pressure is considered an adequate performance criteria for pressurized aircraft, 

while aircraft without cabin pressurization should have static systems with leak rates 

235 
AC 23-17B 04/12/05 

that do not exceed 100 feet per minute at a test pressure equivalent to 1,000 feet 

pressure differential. It was pointed out that the leak test should be considered a 

convenient and simple method of proving the integrity of the static pressure system 

and not a quantitative test of the system error. The Agency agrees with these 

comments and changes to the leak rate requirements of Secs. 23.1325 and 25.1325 

consistent with the recommendations are proposed herein. 

In addition to the foregoing, it is also proposed to change the requirements of Sec. 

23.1325 to require that a correction card be provided if the altimeter indication on 

the alternate system differs by more than 50 feet from the altimeter indication on the 

normal system. The present requirement permits the use of a correction card if the 

reading of an altimeter on the alternate static pressure system exceeds a 2-percent 

tolerance. The Agency considers that this latter value is too small at low altitudes 

and too large at high altitudes.” 

The proposed rule was changed by Final Rule, Docket 7831, as follows: “One 

commentator pointed out that the requirement that the proof tests of the static 

pressure systems on unpressurized aircraft be conducted with the static pressure 

system evacuated to a pressure differential based on an altimeter reading of 1,000 

feet at sea level could be confusing to people located at elevations above 1,000 feet 

mean sea level. The FAA agrees and Secs. 23.1325(b) (2) (i) and 25.1325 (c) (2) (i)

have been further amended to permit the proof test of unpressurized aircraft with the 

static pressure system evacuated to an altimeter reading of 1,000 feet above the 

airplane elevation at the time of the test. One commenter stated that the proposed 

Sec. 23.1325(b)(3) requirement for a correction card where altimeter readings on 

primary and alternate static systems differ by more than 50 feet, is too restrictive at 

high altitudes and high Mach numbers and suggested clarification as to the range of 

altitude and Mach numbers applying to this tolerance. However, the FAA believes 

that the proposed requirement is necessary to assure proper vertical separation 

considering the entire altitude-speed range. In the high Mach-high altitude regime, 

the static system accuracy may be marginal at best and errors introduced while on 

the alternate system could lead to hazardous operation if the pilot is not informed of 

the magnitude of the error. In the low speed-low altitude regime, static system errors 

are minimized so that correspondence between the two systems should pose no 

problem.” 

Amendment 23-20 and Subsequent

A proposed revision to NPRM 75-10 explains this amendment as follows: “The 

proposal would provide more complete duality of static pressure sources if both 

primary and alternate sources are installed. Service experience, however, indicates 

no need for an additional requirement that the selector be secured in the primary 

source position.”

236 
04/12/05 AC 23-17B 

Amendment 23-34 and Subsequent

See AC 23-8B, “Flight Test Guide for Certification of Part 23 Airplanes,” for 

additional guidance. 

This amendment added commuter category requirements from “Part 135, 

Appendix A., section 14.” 

Amendment 23-42 

Paragraph (g) was added. For airplanes specifically prohibited from flight in 

instrument meteorological conditions and icing conditions according to 

§ 23.1559(b) of this part, paragraph (b)(3) of this section does not apply. 

This proposal would allow airplanes that are specifically prohibited from flight in 

instrument meteorological conditions and IFR icing conditions to be certificated 

without an alternate static air source. 

A proposed revision to NPRM 89-5 explains Amendment 23-42 as follows: “The 

FAA reviewed the rationale supporting the requirements in Sec. 23.1325(b)(3) that 

"each static pressure port must be designed or located in such a manner that the 

correlation between air pressure in the static pressure system and true ambient 

pressure is not altered when the airplane encounters icing condition," and "an antiicing means or an alternate source of static pressure may be used in showing 

compliance with this requirement." The primary concern was that airframe ice 

accumulation would disturb airflow in the vicinity of static port(s) causing errors in 

the static pressure systems and altimeter indications. 

The need for such a requirement, as stated in Notice 64-14 (29 FR 3310, March 12, 

1964) was based on IFR operations at higher airspeeds and altitudes above 14,500 

feet. The purpose was "to increase safety and improve airspace utilization" (vertical 

separation of air traffic). 

In an airplane certificated for flight in IFR conditions, an applicant can show 

compliance without flying the airplane in icing conditions. For example, if the 

airplane were equipped with a pitot-static probe, anti-icing would be appropriate. If 

the airplane had static pressure ports installed on the fuselage an alternate static 

pressure source would suffice. 

Under the present requirements, an airplane limited to approval for flight in VFR 

conditions must meet a requirement intended to provide better vertical separation for 

airplanes flying in IFR conditions or icing conditions at altitudes above 14,500 feet. 

Section 23.1325(b)(3) requires that all small airplanes, including an airplane, which 

isn't required to have a sensitive altimeter or a heated pitot, must either be tested in 

237 
AC 23-17B 04/12/05 

icing conditions or show compliance by a heated static pressure source or an alternate 

static pressure source. 

Therefore, it is considered inappropriate to continue to impose a requirement for an 

alternate static source or a means for anti-icing the static source on airplanes 

specifically prohibited from flight in IFR or icing conditions. It is considered 

unlikely that a midair collision could be caused by the altitude error in an airplane 

flying VFR that accidentally enters icing conditions. Continued flight in such 

conditions cannot be sustained for long periods of airplanes without anti-icing or 

deicing equipment and most airplanes certificated for VFR only do not fly above 

14,500 feet. Also, an FAA review of incidents of static system malfunction in VFR 

icing conditions for the past five years showed no reported incidents or accidents.” 

Amendment 23-49 and Subsequent 

Static pressure system icing was addressed in Amendment 23-49 as follows: “Current 

Sec. 23.1325(b)(3) establishes certain static pressure system requirements for 

airplanes that encounter icy conditions. Current Sec. 23.1325(g) exempts from the 

requirements of (b)(3) airplanes that are prohibited from flight in instrument 

meteorological conditions in accordance with Sec. 23.1559(b). After the adoption of 

Sec. 23.1325(g), it came to the FAAs attention that there are conditions other than 

instrument meteorological conditions where icing may be encountered and, therefore, 

that this paragraph should also exempt from the provisions of Sec. 23.1325(b)(3) 

airplanes that are prohibited from flight in icing conditions. Accordingly, Sec. 

23.1325(g) would be revised to read, "For airplanes prohibited from flight in 

instrument meteorological or icing conditions." 

Amendment 23-50 and Subsequent 

A proposed revision to NPRM 94-22 explains this amendment as follows: “This 

proposal would revise Sec. 23.1325 (e) to clarify that the calibration must be 

conducted in flight, which is standard practice. The text of Sec. 23.1325 (f) would be 

removed and the paragraph would be reserved. The text of paragraph (g) would be 

moved to paragraph (f) in a future rulemaking action. The results of the calibration 

would be required in the proposed Sec. 23.1587.” 

238 
04/12/05 AC 23-17B 

23.1326 Pitot heat indication systems

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this rule as follows: “Due to 

advancements in technology, many part 23 airplane installations now utilize 

equipment whose data sources are critical to the accurate and dependable operation 

of that equipment. The heated pitot tube is one such data source. The pitot heat 

indicating system will advise the pilots of any inoperative heating element in the pitot 

tube and that subsequent inaccuracies may result. 

Part 23 airplanes certificated for flight under instrument flight rules or for flight in 

icing conditions are required by current Sec. 23.1323(e) to have a heated pitot system 

or an equivalent means of preventing an airspeed indicating system malfunction due 

to ice accumulation. This proposal would require such airplanes equipped with a 

heated pitot tube to be equipped with a pitot tube heat indicating system. This 

requirement will provide greater assurance that the pilots will not be dangerously 

misled by faulty flight instrument indications caused by pitot tube icing. 

When pitot tube heat indicating system requirements were added to part 25, the FAA 

noted the occurrence of at least one accident and several incidents in which an 

airspeed indicating error occurred that might have been avoided if a pitot tube heat 

indicating system had been installed. Part 23 airplanes operate at lower airspeeds 

and over shorter distances that do part 25 airplanes; therefore, their exposure to 

moisture and temperature conditions where icing may occur is higher than it is for 

transport category airplanes. Because of this environmental exposure, the potential 

for an inoperative heated pitot tube becoming a hazard to part 23 airplanes is 

greater.” 

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 because of 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 as follows: 

(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. 

239 
AC 23-17B 04/12/05 

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 ELOS finding that preserves a “dark cockpit” provided a placard or 

flight manual prescribes when to operate the pitot heat. 

240 
04/12/05 AC 23-17B 

23.1327 Magnetic direction indicator 

The corresponding rule in CAR 3 is CAR 3.666. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.11.

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 an appropriate placard to dictate which 

electrical devices should be switched off when reading the magnetic compass, 

according to § 23.1547, equivalent safety under part 21, § 21.21(b)(1) may be shown. 

On 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 (nonstabilized). 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 confirmed 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 airplanes 

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 ELOS 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 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 

ELOS finding. The airplanes 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: 

241 
AC 23-17B 04/12/05 

(1) The accuracy is not excessively affected by the airplanes vibration or 

magnetic fields. 

(2) Deviations of more than 10 degrees in level flight are not permissible, unless 

a magnetic stabilized direction indicator, which does not have a deviation in 

level flight greater than 10 degrees 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 degrees 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 degrees. 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this amendment as follows: “This 

proposal would allow deviation (exceeding 10 degrees) in a magnetic nonstabilized 

direction indicator caused by operation of electrically powered systems if the 

airplane is equipped with another direction indicator that meets the standard 

specified in paragraph (b) and the airplane is placarded in accordance with 

proposed Sec. 23.1547(e). Under these conditions, safety would not be adversely 

affected.” 

The proposed rule, revised by Final Rule, Docket 14625, is explained as follows: 

“One commentator objected to the proposal to amend Section 23.1327 on the ground 

that it would allow an indefinite amount of interference, which would confuse the 

pilot. The FAA believes that the placard that would be required by proposed 

Section 23.1327(c) would serve to alert the pilot to the fact that certain electrical 

loads, when switched on, cause excessive deviations of the magnetic nonstabilized 

direction indicator. The commentator further stated that if an additional magnetic 

direction indicator (having a deviation less than 10 degrees) were installed, there 

would no longer be a need for the magnetic nonstabilized direction indicator. The 

FAA agrees that an additional magnetic nonstabilized direction indicator is not a 

practical alternative, and proposed Section 23.1327 is revised to delete that 

alternative. However, the FAA believes that the proposed exception is appropriate 

for those instances in which a magnetic stabilized direction indicator or gyroscopic 

direction indicator is installed along with a magnetic nonstabilized direction 

indicator. The commentator also contended that the gyroscopic direction indicator 

242 
04/12/05 AC 23-17B 

alternative is not a practical solution because it must be reset frequently, thus 

increasing the pilots workload. The FAA does not believe that the need to 

periodically reset the gyroscopic direction indicator adds significantly to the pilot's 

workload. Finally, this commentator stated that proposed Section 23.1327(c) would 

allow interference that is contrary to current Secs. 23.1327(a) and 23.1301(a)(4). 

With respect to current Section 23.1327(a), the FAA believes the comment is valid, 

and that paragraph is revised for clarification. In view of this change, the 

requirements in current Section 23.1327(a) and proposed Section 23.1327(b) are 

combined into Section 23.1327(a), and proposed Section 23.1327(c) is redesignated 

Section 23.1327(b). With respect to current Section 23.1301(a)(4), the FAA does not 

believe the comment is valid, since the system consisting of a magnetic nonstabilized 

direction indicator and either a magnetic stabilized type or a gyroscopic type would 

meet that requirement.” 

243 
AC 23-17B 04/12/05 

23.1329 Automatic pilot system 

The corresponding rule in CAR 3 is CAR 3.667. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.12.

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: 

(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), in part 23 airplanes, should consider the 

compatibility of the autopilot system with the increased horsepower. This is because 

the malfunction and performance tests of the autopilot are conducted with a defined 

amount of engine power. 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 the original approval of the autopilot system is still valid.

244 
04/12/05 AC 23-17B 

a. The results of malfunction testing determine which flight condition is most 

critical. The effects of autopilot runaways are more pronounced at aft CG. Also, 

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. 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 CG and minimum autopilot authority. This 

configuration is used to show 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 CAR for 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 torques. 

§ 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.

245 
AC 23-17B 04/12/05 

§ 23.1322 Warning, caution, and advisory lights. 

§ 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 “RTCA Document DO-160D”, July 21, 1998. 

AC 23.1309-1C “Equipment, Systems, and Installations in Part 23 

Airplanes”, March 12, 1999. 

c. Technical Standard Order (TSO) 

TSO-C9c “Automatic Pilots”, September 15, 1960. 

d. Industry Documents 

RTCA/DO-160D “Environmental Conditions and Test Procedures 

for Airborne Equipment”, July 29, 1997. 

RTCA/DO-178B “Software Considerations in Airborne Systems 

and Equipment Certification”, December 1, 1993. 

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 cancelled in 

1977, criteria essentially equivalent to that contained in it continued to be used to 

show compliance with the applicable autopilot installation requirements. 

AC 23.1329-1 was canceled by incorporation into this AC. 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: 

246 
04/12/05 AC 23-17B 

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 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. Thus, these forces, as measured at the pilots' 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 five 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 three-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 a reliable failure warning system

in the climb, cruise, and descent flight regimes. 

(2) A one-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, according to 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: 

(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. 

247 
AC 23-17B 04/12/05 

(2) Annunciators should conform to the proper color as specified in § 23.1322. 

(3) A determination 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 used 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 to 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 CG/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 airplanes 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 resulting 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 resulting corrective 

action would result in loads beyond the design limit loads of the 

airplane. 

248 
04/12/05 AC 23-17B 

(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 degrees or pitch attitude in excess of + 30 degrees deviation from

the attitude at which the malfunction was introduced. 

(v) A hazardous dynamic condition. 

(2) Normal Flight Malfunctions 

The airplanes 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. Amendment 

3-2 to part 3 of the CAR, effective August 12, 1957, requires that an 

autopilot system should be designed so a single malfunction will not 

produce a hardover signal in more than one control axis (reference §§ 

3.667(e) and 23.1329(e)). 

NOTE 18: A three-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 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) and (ii) (titled, “Normal 

249 
AC 23-17B 04/12/05 

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 AFM or approved manual 

material. In maneuvering and approach flight regimes, the pilot should 

recognize an autopilot system malfunction within one-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 19: Accurate measurement of altitude loss, because of 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 because of 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 degrees. 

(2) Pitch to exceed +45 degrees, -35 degrees. 

(3) Accelerations outside the 0g to 2.5g envelope. 

250 
04/12/05 AC 23-17B 

(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 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 degrees. 

(2) Pitch to exceed +45 degrees, -35 degrees. 

(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 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. 

251 
AC 23-17B 04/12/05 

(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 the failure effect of a 

combined monitor failure and autopilot is 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 CG 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 three degrees per second or 

produce hazardous attitudes. This rate should be measured and averaged over a 

five-second period. If approval is sought for ILS approaches initiated with one 

engine inoperative, the autopilot should be capable of conducting the approach. 

252 
04/12/05 AC 23-17B 

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 Pilots Operating Handbook (POH) 

or presented to the pilot in the form of placards) as follows: 

(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 preengagement 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, because of possible 

malfunctioning of the autopilot system. 

(ii) A statement of the altitude loss because of 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.) 

253 
AC 23-17B 04/12/05 

FIGURE 1. ACCEPTABLE METHOD FOR DETERMINING 

ALTITUDE LOSS IN APPROACH 

254 
04/12/05 AC 23-17B 

FIGURE 1⎯CONTINUATION 

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 told 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 a reliable failure warning system. After a failure, 

recovery should be initiated one second after the pilot recognizes the failure. 

a. A three-degree glide slope should be used for these tests 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 shown 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. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 added a requirement for a quick release 

explained as follows: “This proposal would standardize the location of the quick 

release (emergency) control for autopilot systems. Standardization permits 

consistency of pilot responses in preventing hazardous airplane attitudes during 

autopilot malfunctions. The location specified is consistent with the requirements for 

part 25 airplanes, except this proposal allows the control to also be located on a stick 

control.” 

255 
AC 23-17B 04/12/05 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 clarified the quick release rule explained as 

follows: “New Sec. 23.1329(b), adopted by Amendment No. 23-24 (58 FR 18958, 

April 9, 1993), does not state clearly that stick controlled airplanes must be equipped 

with the same autopilot quick release controls that are required for airplanes with 

control wheels. This proposed revision of Sec. 23.1329(b) would clarify that a quick 

release control must be installed on each control stick of an airplane that can be 

operated from either pilot seat.”

256 
04/12/05 AC 23-17B 

23.1331 Instruments using a power source 

The corresponding rule in CAR 3 is CAR 3.668. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.14.

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. 

NOTE 20: Section 23.1331 does not apply to pitot and static systems. 

Electrical Systems 

When complying with paragraph (b) in this section, a single battery required for 

starting is acceptable if: 1) the electrical system is capable of continuous normal 

operation without external excitation or stability, and 2) 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 (that is, loss of voltage regulation). 

This would not be corrected 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 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. 

257
AC 23-17B 04/12/05 

The standby system should either supply enough 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 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 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 were not required, the standby installation would be for 

non-required equipment according to 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 assure 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 AFMS or placards, as appropriate. This 

information should meet the requirements of §§ 23.1583 and 23.1585 of this part, and 

it should emphasize the standby vacuum system is for emergency use only and should 

not be used for dispatch purposes. 

Amendment 23-43 and Subsequent

Independent Power Sources 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal requires a visual annunciation to indicate when power for gyroscopic 

instruments is not adequate, and two independent sources of power for all airplanes. 

Requirements in current paragraphs (a)(1) and (a)(2) are being deleted because the 

general requirements of §§ 23.1301 and 23.1309 will adequately address these 

issues.” 

The proposal to Amendment 23-43 by Final Rule, Docket 26344, is based on 

comments as follows: “Several comments were received on this proposal. One 

commenter supports the proposal but notes that it does not address non-gyroscopic 

instruments, and would result in power supply requirements for such instruments 

being omitted from regulations. This commenter also believes the word "adjacent" in 

proposed Sec. 23.1331(a) is too restrictive and requests the meaning of the word 

"independent" in proposed Sec. 23.1331(c) in context of sources of power for singleengine airplanes. Finally, this commenter identifies support for the provisions of 

proposed Sec. 23.1331(b)(2).

258
04/12/05 AC 23-17B 

The FAA has reviewed this proposal and agrees that by inserting the word 

"gyroscopic" in the introductory text of this proposal, other types of instruments that 

use a power source for their function would be omitted from the regulations. To 

correct this omission and retain the current provision of the regulations that address 

all instruments, the word "gyroscopic" is being removed from the introductory text. 

To further clarify the applicability of these requirements, the words "that uses a 

power source" are being added between the words "instrument" and "the". 

The FAA also reviewed this commenters position on the word "adjacent" and agrees 

that its application could be too restrictive. The intent of this proposal is to require 

any installed separate power indicator to be located so that a pilot who is using that 

instrument will notice the loss of that instrument's power. To clarify and preclude 

restrictive application of this requirement, the word "adjacent" is removed and 

replaced with words similar to those used in Sec. 23.1321(a). 

In regard to the commenters question on the word "independent" in context of 

sources of power for single-engine airplanes, this word has the same meaning for all 

airplanes, except that on single-engine airplanes the second source cannot be driven 

by a separate engine. 

Adverse service experience that has resulted from power source failures and the 

subsequent loss of flight instrumentation has shown that it is necessary to provide a 

backup power source for the flight instruments. In the case of instruments that use a 

vacuum power source, the second source has been provided by installing a smaller 

electric driven vacuum pump and by arranging the vacuum system so that this pump 

is isolated from the normal vacuum system and so that it provides power to the 

instruments only after the normal engine driven pump fails.” 

This amendment adds the requirement for independent power sources for required 

instruments for single-engine as well as multiengine airplanes. This was considered 

appropriate because of 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 so any failure of one power 

source or instrument will not cause the complete loss of a required flight 

instrument function. Thus, in the 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. 

259
AC 23-17B 04/12/05 

b. Ships 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. 

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. Thus, 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 enough 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 installing dual alternators or vacuum systems 

on single engine airplanes. Other options include: 1) a dedicated battery with a 

30-minute capacity for electrical instrument loads essential to continued safe 

flight and landing, 2) use of differently powered types of instruments for primary 

and standby, or 3) 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 so far); and 

(2) Require the state of charge be available to the pilot during all stages of flight 

and include a caution alert according to § 23.1322 when the state of charge 

is less than 30 minutes. 

260
04/12/05 AC 23-17B 

23.1335 Flight director systems

The corresponding rule in CAR 3 is CAR 3.669. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.15. 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explained the commonality between parts 23 

and 25 as follows: “Service experience has shown that safety is compromised when 

the flight crew is not informed as to the current mode of operation of the flight 

director system. Moreover, experience has also shown that the position of the mode 

selector switch is not a reliable means of providing that information. This proposal 

would ensure that an indication entirely independent of the mode selector switch is 

provided.” 

The proposals for both parts 23 and 25 for Amendment 23-20, of Final Rule, Docket 

14625, is based on the following comment: “One commentator criticized the 

proposed phrase "independent of the mode selector switch" because it would tend to 

dictate design. Another commentator objected to the same phrase on the ground that 

it does not take into account modern panels, which incorporate illuminated mode 

indications that give positive indication of the selected mode, but are not independent 

of the mode selector switch since they are incorporated in it. The FAA agrees with 

these comments. The phrase "independent of the mode selector switch" is deleted 

from proposed Section 25.1335 and another sentence is added thereto reading, 

"Selector switch position is not acceptable as a means of indication." 

261
AC 23-17B 04/12/05 

23.1337 Powerplant instruments installation 

The corresponding rules in CAR 3 are CAR 3.670, 3.671, 3.672, 3.673, and 3.674. 

The corresponding rules in the Airship Design Criteria, FAA-P-8110-2, Change 2, are 

sections 6.4 and 6.16.

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 (according to 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 23-16A, “Powerplant Guide for Certification of Part 23 Airplanes and 

Airships,” § 23.959, for unusable fuel test procedures for guidance on 

§ 23.1337(b)(1). AC 23-16 incorporated the guidance in canceled AC 23.959-1, 

“Unusable Fuel Test Procedures for Small Airplanes.” The quantity of unusable fuel 

is determined by compliance to § 23.959, and AC 23-16A provides an acceptable 

means of compliance. 

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 (Reciprocating 

Engine Aircraft),” or MIL-G-9798. 

The rule requires that fuel quantity be calibrated as zero when only unusable fuel (as 

determined under § 23.959) 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 ELOS 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 fuel quantity indicator. 

However, when installed in a reciprocating engine airplane and produced under TSOC55 authorization, the allowable error of the indicator is no more than three percent 

of full scale. It is believed that ground and cruise attitude(s) are the minimum

attitudes required for operation within the three 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. 

262
04/12/05 AC 23-17B 

The purpose of this evaluation is to establish the ground and cruise attitude(s) gauge 

tolerance is functional and satisfactory for the airplane. 

Amendment 23-7 and Subsequent

This amendment moved the requirement for cylinder head temperature indication 

from this section to § 23.1305. 

Amendment 23-18 and Subsequent

See AC 23-8B, “Flight Test Guide for Certification of Part 23 Airplanes,” for 

guidance on fuel quantity indicators and auxiliary tanks. 

A proposed revision to NPRM 75-19 revised paragraph (a) and is explained as 

follows: “Powerplant instruments and instrument lines sometimes utilize flammable 

fluids. The proposal would require that those instruments and lines be installed and 

located so that leakage of the fluid would not create a hazard. The proposal would 

also editorially revise Sec. 23.1337(a) for clarity.”

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 adds auxiliary power unit requirements and 

clarifies a fuel quantity rule in paragraph (b)(5) as follows: “This proposal adds APU 

installation requirements and clarifies fuel quantity indicator requirements. 

Conference Proposal 449 recommended amending this section to include APU 

requirements for the reason that applications have been received for approval of 

auxiliary power unit installations in part 23 airplanes. These installations also need 

protection from the escape of flammable fluids. The proposed revision to paragraph 

(b)(5) will maintain compatibility with proposed changes to Section 23.955(d).” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this amendment as follows: “This 

proposal would revise the heading of this section to reflect the powerplant instrument 

installation requirements that it contains. The difference between this section and 

Sec. 23.1305 is clarified by this change. 

Section 23.1337(b) would be revised by removing the wording that authorizes 

installation of only those fuel indicators marked in gallons and pounds. In countries 

that use the metric system, other acceptable units of measure for marking fuel 

indicators are used. This proposed revision would allow the use of any appropriate 

measurement unit. 

263
AC 23-17B 04/12/05 

Section 23.1337(b) would also be revised by adding the word "usable" to the first 

sentence of this section. This revision is consistent with the requirements of Sec. 

23.1337(b)(1), which requires the fuel quantity indicator to be calibrated to read 

"zero" when the fuel in the tank is equal to the unusable fuel determined under 

Sec. 23.959. 

Proposed new Sec. 23.1337(b)(4) would require a "means to indicate" the amount of 

usable fuel in each tank when the airplane is on the ground. This requirement would 

ensure that a reliable means is provided for the pilot to determine before takeoff that 

the amount of fuel that is in the airplane is adequate for the intended flight. The 

ability to make this preflight determination will help reduce the number of accidents 

that have resulted from fuel starvation. This proposal, which is patterned after 

Sec. 23.1337(d) and (d)(1), would not require a separate fuel indicating system. The 

means to determine the amount of fuel while on the ground may be provided by a 

calibrated dipstick separate markings on the inflight fuel indicator, or any other 

acceptable means selected by the manufacturer. Accordingly, this proposal would 

contribute to the safe operation of the airplane and would not appreciably add to the 

cost of the airplane design.” 

Amendment 23-51 and Subsequent 

A proposed revision to NPRM 94-19 clarifies the reference to Subpart E as follows: 

“Under the area of "Installation," the reference in Sec. 23.1337 (b)(1) to Sec. 23.959 

would be changed to Sec. 23.959 (a), in accordance with the revision to Sec. 23.959 

proposed in this notice. The revision would redesignate the existing Sec. 23.959 text 

as Sec. 23.959 (a); there is no change in the requirement itself.”

264
04/12/05 AC 23-17B 

ELECTRICAL SYSTEMS AND EQUIPMENT 

23.1351 General 

The corresponding rules in CAR 3 are CAR 3.681, 3.682, 3.686, and 3.687. 

The corresponding rules in the Airship Design Criteria, FAA-P-8110-2, Change 2, are 

sections 6.4 and 6.17.

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 the advantages of having a battery connected for initial excitation or 

stabilization for alternators needing it outweighed the results of that battery failing. 

This exception is only applicable to alternator installations that need a battery. Loss 

of an alternator because of battery failure was considered of no greater result than the 

intrinsic failure of the alternator itself. In 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 (that is, shorts to ground and conductor failure) that 

would also result in loss of the electrical system. 

Wire, meeting MIL-W-5086, has been removed from the list of approved wire in AC 

43.13-1B, Change 1, because of its flammability characteristics, corrosive vapors, and 

toxic gases of Polyvinyl Chloride (PVC) insulation. See AC 43.13-1B, Change 1, 

“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, Change 1, should be shown by tests and 

analyses to meet the airworthiness requirements of § 23.863, 23.1351, 23.1359, and 

23.1365 as applicable. 

Amendment 23-14 and Subsequent 

This amendment added requirements so designed the risk of electrical shock to crew, 

passengers, and ground personnel is reduced to a minimum. There must also be a 

means to give immediate warning to the flight crew of a failure of any generator. 

265
AC 23-17B 04/12/05 

Fire resistance. Electrical equipment must be so designed and installed that if there 

is a fire in the engine compartment, during which the surface of the firewall adjacent 

to the fire is heated to 2000 degrees F for five minutes or to a lesser temperature 

confirmed by the applicant, the equipment essential to continued safe operation and 

located behind the firewall will function satisfactorily and will not create an 

additional fire hazard. 

Amendment 23-17 and Subsequent 

A proposed revision to NPRM 75-10 requires overvoltage control of generators with 

the following explanation: “Complete electrical system failures on both single and 

twin engine aircraft continue to occur. The electrical failure that causes this problem 

is often the loss of voltage control in the voltage regulator. This usually results in a 

bus voltage well above the capabilities of the electrical equipment connected to the 

bus. This overvoltage condition frequently destroys electronic equipment and boils 

the electrolyte in the battery. The proposal is only directed at this overvoltage 

problem and therefore differs from Sec. 23.1351(b). The proposal would not require 

a specific method of overvoltage control.” 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 added external power requirements with the 

following explanation: “If an external power supply with reverse polarity or reverse 

phase sequence were to supply power to the airplanes electrical system extensive 

damage to the system could result. This proposal would require a means to prevent 

such an occurrence.” 

Amendment 23-34 and Subsequent 

A proposed revision to NPRM 83-17 added commuter category requirements from

“part 135, Appendix A, sections 61 and 63”.

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal would allow a generator to be installed and operate below its continuous 

rating when it has a rating higher than necessary, allow methods other than reverse 

current cutouts for protecting against reverse current, and would require the airplane 

to operate safely for 5 minutes without normal electrical power. 

Conference proposal 453 recommended revising paragraphs (c)(1) and (c)(3)

essentially as proposed herein. The justification given is that the generator-rated 

output may be higher than required for the electrical loads of the airplane and, in 

such case, the electrical system (generator output) is limited by its regulation system. 

266
04/12/05 AC 23-17B 

The consensus at the conference supported the objective of this conference proposal. 

The FAA agrees with the need for changing paragraph (c)(1) to clarify its intent and 

to revise paragraph (c)(3) to relieve the burden to install a specific type of reverse 

current control where more efficient and less costly controls are now available. 

Conference proposal 456 recommended adding paragraph (g), essentially as 

proposed herein except the requirement would only be applicable for airplanes 

operated above 25,000 feet. The justification was that part 23 airplanes that operate 

at high altitudes above 25,000 feet depend upon electrical power for safe operation. 

Emergencies involving loss of normal electrical power at or above this altitude 

typically result in the loss of other systems, such as electric fuel pumps, 

pressurization system, warning system, navigation, communications, and 

instrumentation. The FAA developed special conditions for part 25 that initiated the 

requirement in this proposal and it was later adopted into part 25 by amendment 

25.41, in 1977. Conference proposal 456 was essentially developed from the part 25 

requirements except for the 25,000 foot applicability. When offered for comment at 

the conference, there were no objections on conference proposal 456. After further 

review, FAA has concluded that the proposal should not be limited to airplanes that 

operate above 25,000 feet since emergencies resulting in the loss of normal electrical 

power are critical for all airplanes. Five minutes is considered adequate time to 

cope with such an emergency so that pilot can operate the airplane safely and assess 

the reason for the loss of normal electrical power. 

Conference proposal 452 recommended changing the phrase "essential for safe 

operation" to "essential to flight safety" for consistency in the regulations. A word 

search of the regulations indicated that there were other phrases such as "essential to 

safety of flight," and "essential to continued safe operation." All of these phrases 

have been interpreted to have the same meaning. Since the affected regulations have 

been administered effectively without significant problems, the FAA does not 

consider the recommended changes to be beneficial.” 

Final Rule, Docket 26324, has the following clarification: “It has been brought to 

the FAAs attention that many electrical generating devices that are used on part 23 

are now referred to as "alternators" and that there is some confusion about such 

units acceptability because Sec. 23.1351(c) continues to address "generators." To 

provide clarifications, ten locations in Sec. 23.1351, paragraphs (c), (c)(1), (c)(2), 

(c)(3), (c)(4), and (c)(5), are being revised by changing the word "generator" to 

"generator/alternator." This proposal is adopted with the aforementioned change.”

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this amendment as follows: “The 

proposal would revise current Sec. 23.1351 by removing portions of paragraphs 

(b)(2) and (b)(3) and by removing all of paragraph (b)(4). The removed 

requirements 

267
AC 23-17B 04/12/05 

are applicable to alternators that depend upon the battery for initial excitation or for 

stabilization. This revision responds to a Joint Aviation Authority recommendation to 

remove the provisions that allow a battery failure to result in the loss of the 

alternator. Information in this recommendation showed that self-excited alternators 

are now available for installation on newly certificated airplanes. The FAA has 

verified that self-excited alternators are now available; therefore, there is no longer a 

need for the regulations to address alternators that depend upon a battery for initial 

excitation and stabilization. 

Revised Sec. 23.1351(c)(3) would require an automatic means for reverse current 

protection. Reverse current protection is accomplished by means that automatically 

detect changes in the current. The proposed revised wording would more accurately 

define this function and the equipment that would accomplish the protection. 

Finally, Sec. 23.1351(f) would be revised by adding a requirement that would require 

the ground power receptacle to be located where its use will not result in a hazard to 

the airplane or to people on the ground using the receptacle.” 

Final Rule, Docket 27806, has the following clarification: “The proposals are 

adopted as proposed, except that paragraph (c)(3) has been revised to clarify that 

protection for any generator/alternator and the airplane electrical system must be 

provided.”

EASA AMC 23.1351(a)(2) is acceptable for FAA certification. 

EASA AMC 23.1351(b)(5)(iv) is acceptable for FAA certification to 14 CFR, part 

23, § 23.1351(b)(4)(iv). 

268
04/12/05 AC 23-17B 

23.1353 Storage battery design and installation 

The corresponding rule in CAR 3 is CAR 3.683. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.18.

Original Issue and Subsequent

See AC 43.13-1B, Change 1, “Acceptable Methods, Techniques, and 

Practices⎯Aircraft Inspection and Repair,” Section 8, for battery installation 

guidance. Replacement batteries would require PMA, unless exempted by 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 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this amendment as follows: 

“Proposed Sec. 23.1353(b)(1) would add "or power" to the battery recharging 

standard. The proposal by adding an alternative standard is intended to clarify the 

present rule to more appropriately include all battery charging means and 

parameters. In some battery charging techniques, voltage is not a controlled factor 

except by response of the battery. Also, power is a measure of heat input to the 

battery. A similar proposal is made for Secs. 25.1353(c)(1)(i), 27.1353(b)(1), and 

29.1353(c)(1)(i). 

Short circuits of nickel cadmium batteries (either internally or to the airplane 

structure through the battery caps) have occurred in service. The heat generated by 

such short circuits may in some circumstances damage nearby structure or essential 

systems. It is therefore proposed to require provisions to prevent any hazardous 

effect from this cause. See also proposed new Sec. 23.1353(f) in Airworthiness 

Review Notice No. 2 (Notice 75-10), which is aimed at reducing the occurrence of 

short circuits that are caused by excessive battery charging current.” 

Amendment 23-21 and Subsequent 

A proposed revision to NPRM 75-10 explains paragraph (g) as follows: “The 

proposal conforms to paragraph (b) of AD 72-19-4, which applies to all nickelcadmium batteries that are capable of being used to start an aircraft engine or 

auxiliary power unit.” 

269
AC 23-17B 04/12/05 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains the 30-minute rule in this amendment 

as follows: “Proposed new Sec. 23.1353(h) would require that, the event of a 

complete loss of the primary electrical power generating system, airplane battery 

capacity must be sufficient to supply at least 30 minutes of electrical power to those 

loads essential to the continued safe flight and landing of the airplane. 

This proposal is not limited to airplanes that are approved for any particular type of 

operation. Although the battery capacity needed for an airplane approved for day 

VFR operations would be much less than the capacity for an airplane approved for 

day/night IFR operations, the same level of safety should be provided for all 

airplanes. While this proposal would add an additional requirement to part 23 for 

normal, utility, acrobatic, and commuter category airplanes, in practice this 

requirement to provide a battery capacity sufficient to supply at least 30 minutes of 

electrical power is not new to many airplane manufacturers. Certain other countries 

in which part 23 airplanes have been certificated have requirements for such a 30-

minute battery capacity. Manufacturers experience with these requirements has 

shown that the only design impact that results from complying with these 

requirements is the need to install a battery with greater capacity than might 

otherwise be installed. Experience has also shown that a load shedding procedure 

may be necessary for certain airplanes. No other airplane design changes would be 

needed. 

Despite the above referenced experience record, this requirement would be new to 

some manufacturers and they may have questions on how it would be applied. For 

that reason, this notice discusses compliance considerations that have emerged from 

experience based on substantively equivalent requirements. 

This compliance experience has shown that the rating of the battery selected for the 

airplane should be sufficient to cover the loss of capacity that would occur with 

battery age and the reduced capacity that results from a realistic state of charge, 

which may be less than a full charge. Using a design battery capacity that is only 75 

percent of the battery nameplate rating would be an acceptable way of accounting for 

these losses. 

In addition to determining the battery rating that would be needed, the manufacturer 

would also need to determine the functions that would be necessary for 30 minutes of 

safe flight and the landing of the airplane. Again, experience has identified several 

functions. For a day VFR approved airplane, no functions may require battery 

power; however, it may be necessary to supply power for certain communication 

capacities or, if the airplane has electrically powered retractable landing gear. 

Providing a secondary means for lowering the gear would be an acceptable 

alternative to providing electrical power or battery power for this function. 

270
04/12/05 AC 23-17B 

For other types of operating approvals, providing power for the following functions 

and equipment should be considered: 

1. Any required flight and navigation instruments. Air driven instruments that would 

function over the required period can also be accepted for this function. 

2. Cockpit and instrument lighting. 

3. For IFR and icing approvals, power for the heated pitot tube. 

4. For radio communication, usually one VHF communication system with power for 

three to five minutes of transmission would be acceptable. 

5. Functions needed for safe night flight and night landing for the airplane. 

6. Electronic engine ignition systems. 

7. Any functions that cannot be readily shed following the loss of generator power. 

8. Engine inlet heat or deicing protection required for normal operation of the 

airplane. 

Although power for the listed functions may provide for the safe operation and 

landing of most airplanes, individual airplane designs may require the consideration 

of additional functions. 

In applying these rules it may be assumed that airframe and engine icing protection 

equipment would not be operating at the time of the generator system failure. Power 

for icing protection would not be required if the icing protection equipment is not 

required for the normal operation of the airplane. 

This proposal would require additional battery capacity and would not alter or 

supersede any other requirements in this part for separate or dedicated emergency 

power supplies. When requirements such as those in current Sec. 23.1331(a) or in 

proposed Sec. 23.1311(a)(5) are applicable to the airplane design, these power 

supplies are required to provide a needed level of safety for that function; therefore, 

that power source must be supplied.” 

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.” 

Previous guidance on this issue states the airplanes primary electrical power includes 

the airplanes electrical generation system and the airplanes starter battery when only 

one battery is installed. The battery for the 30-minute criteria, therefore, should be an 

independent power source from the airplanes starter battery. If adequate monitoring 

and procedures are incorporated so the pilot knows the airplanes starter battery 

271
AC 23-17B 04/12/05 

meets the 30-minute criteria after an engine start and during all other operations, an 

ELOS finding may be an acceptable method for using the airplanes 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: 

“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. Sometimes, 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 either 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 safety gained 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 parts 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

272
04/12/05 AC 23-17B 

• 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 as 

follows: 

(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. 

273
AC 23-17B 04/12/05 

23.1357 Circuit protective devices 

The corresponding rules in CAR 3 are CAR 3.690, 3.691 and 3.692. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.19.

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 ACO, 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) on 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 and high oil temperature, to mitigate the 

effects of loss of indication. 

274
04/12/05 AC 23-17B 

For part 23 applications, the definitions of a switch and a circuit breaker are as 

follows: 1) A switch is a device for opening and closing or for changing the 

connection of a circuit; 2) 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. Thus, circuit breakers used for operational functions are not 

acceptable because 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. 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains § 23.1357(b) as follows: “Proposed 

paragraph (b) would clarify the present rule by specifically prohibiting the use of a 

single protective device to protect an essential and a nonessential circuit, or more 

than one essential circuit. A protective device protecting two circuits would trip in 

response to a fault in either one; thus a fault in the nonessential circuit would render 

the essential circuit inoperative.” 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “The 

intent has historically been to protect the airplane from the hazards of all electrical 

faults. Paragraph (a)(1) exempted starting motor circuits because they did not have 

power applied, except during engine starting. It is proposed to clarify the intent of 

paragraph (a)(1). The existing rules of paragraph (e) require spare fuses for all 

electric circuits. This proposal would require spare fuses for fuses identified as 

replaceable in flight, which would be those required by paragraph (d) and any other 

fuses identified as replaceable in flight. This proposal would also require the fuses 

be readily accessible and available.” 

275
AC 23-17B 04/12/05 

23.1359 Electrical system fire protection

See guidance in section 23.853 of this AC. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this rule as follows: “Proposed new 

Sec. 23.1359 would require smoke and fire protection for electrical system 

installations. The provisions of Sec. 23.1359(a) of this proposal state that electrical 

systems must meet the applicable requirements of Secs. 23.863 and 23.1182. 

Proposed Sec. 23.1359(b) would require that the electrical systems components 

installed in designated fire zones and used during emergency procedures be fire 

resistant. This provision is needed to clarify the requirements for electrical system 

components that may be installed in the designated fire zones identified in Sec. 

23.1181. 

Finally, Sec. 23.1359(c) provides burn criteria for electrical wire and cables. A 

proposed revision to appendix F of part 23 that would add appropriate wire testing 

criteria is included in this notice. 

This proposed burn criteria for wire is necessary because of the increased use of 

electrical systems in the design of part 23 airplanes and the resulting increase in the 

amount of electrical wire being installed. This increased use results in the need to 

ensure that wire insulating material does not become the source of an in-flight fire 

and/or that it does not propagate a fire from another source. The electrical wire 

burn requirements in this proposal, along with the testing identified in revised 

appendix F, would ensure that installed electrical wire has insulating material that 

reduces the possibility of hazardous in-flight fires.” 

276
04/12/05 AC 23-17B 

23.1361 Master switch arrangement

The corresponding rule in CAR 3 is CAR 3.688. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.20. 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this amendment as follows: “Circuits 

by-passing the master switch remain energized after the master switch is opened. If 

such circuits fail during a crash landing they might possibly ignite nearby flammable 

fluids or vapors. This proposal would reduce that possibility.” 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “The 

proposal clarifies the master switch arrangement requirement and permits new 

generations of engines to operate with the master switch turned off, as is necessary to 

isolate hazardous electrical faults. 

Subsequent to the receipt of the proposals submitted for the Part 23 Review, engine 

designs have been developed that depend on an electrical power source for normal 

ignition and/or fuel pressure. An electrical fault that makes it necessary to turn off 

the master switches must not cause unintentional disabling of such designed engines. 

However, the pilot must retain the capability to isolate all sources of electrical 

energy that might ignite flammable fluids that are likely to escape during a survivable 

crash landing. 

When conference proposal 461 was discussed, a commenter specifically noted that he 

had no objections to the multiple circuits restriction in the proposal and supported 

that position. The commenter did express a concern that the proposal, "Load, 

circuits, such as cabin entry lights whose functions are needed prior to entering the 

cockpit," may unnecessarily limit those circuits to those functions that are needed 

before entering the cockpit. It was noted that there are other continuously energized 

circuits that do meet the "needed prior to entering the cockpit" definition in the 

proposal. A circuit for an electrical clock was cited as an example. 

The FAA has reviewed the proposal and this discussion and agrees that while the 

proposed wordage was only intended to provide an example of the type of circuit that 

was permitted by this section, it could be interpreted as more restrictive than 

intended; therefore, this example language has not been included in this proposal. 

This action by FAA should not be interpreted as an endorsement to install an 

unlimited number of circuits that bypass the master switch. This provision was added 

277
AC 23-17B 04/12/05 

to, and retained in, the requirements because it was recognized that there are a 

limited number of electrical functions that are needed when the master switch is in 

the position. The requirements of this section provide for the safe installation of 

these circuits. The five-ampere load restriction of a new paragraph (b)(3) was added 

because the FAA was made aware of an installation in which this provision was 

being used to circumvent the master switch arrangements by using up to four fiveampere fuses to supply a 20-ampere circuit. This restrictive provision should make it 

clear that such installations are not permitted.” 

The proposal in NPRM 90-23 was revised by Final Rule, Docket 26344 as follows: 

“This proposes to amend Sec. 23.1361 to clarify the requirement for the master 

switch arrangement and to permit new generations of engines to operate with the 

master switch turned off. No comments were received; however, an editorial revision 

has been made that revises the text of the last sentence from one that permits the 

master switch arrangement to use separate switches to text that provides 

requirements for the master switch arrangement if separate switches are installed. 

This proposal is adopted with the aforementioned change.” 

Amendment 23-49 and Subsequent 

This amendment is explained by Final Rule, Docket 29806 as follows: “To 

harmonize with the JAR this proposal would revise Sec. 23.1361(c) by making an 

editorial change to remove the last two words of the paragraph that read "in flight." 

This change will not alter the meaning of the requirement.” 

278
04/12/05 AC 23-17B 

23.1365 Electric cables and equipment 

The corresponding rule in CAR 3 is CAR 3.693. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.21.

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, 

Change 1, “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. A proposed revision to 

NPRM 71-13 explains this amendment as follows: “This proposal is considered 

necessary because of the higher powered electrical systems being installed in part 23 

airplanes. Such systems are more likely to cause heat damage and the emission of 

toxic fumes in the event of malfunction.” 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal provides crashworthiness standards for electrical cables. It would require 

that electrical cables be designed to allow a reasonable degree of deformation and 

stretching without failure and be isolated from flammable fluid lines or must be 

shrouded in insulated flexible conduit.” 

The proposal in NPRM 90-23 was changed by Final Rule, Docket 26344, and is 

explained as follows: “One commenter believes the word "isolated" used in 

proposed Sec. 23.1365(c)(1) is not compatible with current practices and suggests the 

word "separated" be used in its place. The FAA reviewed this recommended change 

and agrees that "separate" better describes the current practice of keeping electrical 

cables and flammable fuel lines spaced a part. This proposal is adopted with the 

aforementioned change.” 

279
AC 23-17B 04/12/05 

Amendment 23-49 and Subsequent 

The proposed revision to NPRM 94-21 has the following explanation for this 

amendment: “This proposal would revise Sec. 23.1365(b) and would add three new 

paragraphs. 

Section 23.1365(b) would be revised in relation to proposed new Sec. 23.1359(c), 

which would require self-extinguishing insulated electrical wires and cables. 

Current Sec. 23.1365(b) requires that cable and associated equipment that would 

overheat in the event of circuit overload or fault must be flame resistant and may not 

emit dangerous quantities of toxic fumes. The proposed revisions to Sec. 23.1365(b)

would remove the reference to electrical cables from the flame resistance 

requirement since the cables would be required to have self-extinguishing insulation 

under Sec. 23.1359(c). The requirement for electrical cables and the associated 

equipment that would overheat to not emit dangerous quantities of toxic fumes has 

been retained. 

The text of Sec. 23.1365(b) that includes the words "at least flame resistant" would 

also be revised by removing the words "at least". The removed words implied that 

there were burn requirements, other than the ones in this section that must be met. 

The three paragraphs that would be added by this proposal would require: (1) The 

identification of electrical cables, terminals, and connectors; (2) the protection of 

electrical cables from damage by external sources; and (3) installation criteria for 

cables that cannot be protected by a circuit protection device. 

As identified in the discussion of proposed Sec. 23.1359, there is an increasing use of 

electrical systems in part 23 airplanes. The resulting increase in the number of 

electrical wires used in part 23 airplanes makes proper installation difficult. The 

proposal for electrical cable identification would provide better assurance that the 

cables will be correctly installed initially and correctly reinstalled when airplane 

maintenance or modifications are accomplished. The other proposed new 

requirements would provide installation criteria that will ensure the protection of 

cables under circumstances that can be expected from the increased use of electrical 

systems.”

280
04/12/05 AC 23-17B 

23.1367 Switches 

The corresponding rule in CAR 3 is CAR 3.694. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.22.

Original Issue and Subsequent

This rule was adopted on February 1, 1965 as a recodification of CAR 3.694 and 

3.695. 

Switches are required to be labeled for 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 

crewmember if manual operation is necessary for safety of flight. 

281
AC 23-17B 04/12/05 

LIGHTS 

23.1381 Instrument lights 

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.696 and 

3.697. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.23. 

282
04/12/05 AC 23-17B 

23.1383 Taxi and landing lights

The corresponding rules in CAR 3 are CAR 3.698 and 3.699. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.24. 

Amendment 23-49 and Subsequent 

This amendment was revised to add taxi lights to the original rule. A proposed 

revision to NPRM 94-21 explains this amendment as follows: “The landing light 

requirements of Sec. 23.1383 would be revised by adding taxi lights to this section. 

When the landing light requirements were included in the normal, utility, acrobatic, 

and commuter category requirements, the same lights were used for both night 

landing and taxiing of the airplane. Due to availability of different types of lights, 

separate lights are now frequently installed for landing and for taxiing. Including the 

word "taxi" in the heading would clarify that the requirements cover both kinds of 

lights. 

Current Sec. 23.1383(a), which requires the lights to be acceptable, would be deleted 

because it is unnecessary to state this. All lights that are found to meet the 

requirements of this section and other directly related airworthiness requirements 

are acceptable. The paragraphs would be redesignated accordingly. 

Current Sec. 23.1383(b)(3) requires that a landing light must be installed to provide 

enough light for a night landing. Proposed Sec. 23.1383(c) would revise "night 

landing" to "night operation" since the requirements would also cover taxiing and 

parking. Proposed new paragraph (d) would require the lights to be installed so that 

they do not cause a fire hazard. This clarifies the need for such an evaluation.” 

283
AC 23-17B 04/12/05 

23.1385 Position light system installation 

The corresponding rule in CAR 3 is CAR 3.700. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.25.

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 

according to §§ 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 

AFM (§ 23.1583(h)). 

Amendment 23-17 and Subsequent 

A proposed revision to NPRM 75-10 explains this amendment as follows: “By 

amending Sec. 23.1385 in a manner substantively identical to that proposed for Sec. 

25.1385. The proposal would revise Sec. 25.1385 to make more clear the 

circumstances in which forward and rear position lights may be located at other than 

the appropriate airplane extremities.” 

The proposed amendment was revised by Final Rule, Docket 14324, and is explained 

as follows: “Two commentators, also suggested that proposed Secs. 23.1385(c) and 

25.1385(c) be revised to permit a new position light to be installed on each wing tip. 

The FAA agrees that further study is necessary to develop factors of general 

applicability for position lights on all aircraft but that a rear position light as far aft 

284
04/12/05 AC 23-17B 

as practical on each wing tip of an airplane is a reasonable alternative location. 

Accordingly, proposed Secs. 23.1385(c) and 25.1385(c) have been revised.”

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal clarifies the location requirements for the position lights, deletes the 

requirement for a single circuit, and removes the redundant statement "must be 

approved". Conference proposal 464 recommended changing paragraph (b) so the 

location of position lights can be compatible with airplane configurations such as 

tandem wing, canards, and swept wings. The words "forward on the airplane" in 

paragraph (b) have been interpreted to mean the first 50 percent of the airplane 

length. 

Conference proposal 465 recommended deleting paragraph (d) since it had been 

interpreted to prohibit multiple circuits from being installed. A consensus at the 

conference supported both of these recommendations. 

The FAA has further studied these issues and concludes clarification is required and 

that the proposed requirements are substantively equivalent to the current rule.” 

285
AC 23-17B 04/12/05 

23.1387 Position light system dihedral angles

The corresponding rule in CAR 3 is CAR 3.701. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.26. 

Amendment 23-12 and Subsequent 

This amendment is explained by Final Rule, Docket 11479 as follows: “Sections 

23.1387, 25.1387, 27.1387, and 29.1387 presently require, in part, that the rear 

position light show unbroken light within a dihedral angle formed by two intersecting 

vertical planes making angles of 70° to the right and left, respectively, of a vertical 

plane passing through the longitudinal axis. Other related provisions of the 

regulations require that the rear position light be mounted as far aft as practicable. 

In certain aircraft designs incorporating swept vertical tail surfaces, the obstructed 

visibility requirements may be met only by locating the rear position light on the 

trailing edge of the rudder. Because this location may cause a number of problems, 

including complex electrical installation and adverse rudder flutter characteristics, 

some manufacturers consider the aft most tip of the fuselage to be a more suitable 

location. Thus, while the rudder position may be farther aft, the fuselage location is 

as far aft as is practicable. At the same time, however, the fuselage location does not 

comply with the obstructed visibility requirements where parts of the rudder and 

vertical stabilizer of a swept tail project into space required to be unobstructed. 

For aircraft having this problem, the obstruction resulting from use of the aft 

fuselage location would, nevertheless, be relatively small because of the thinness of 

the vertical stabilizer and rudder. Moreover, the obstruction occurs at a high angle 

above the longitudinal axis of the aircraft so that except for the near-zenith position, 

the rear position light shows unbroken light. 

Related requirements for position lights allow diminishing light intensity with 

increasing angle above or below the horizontal. Thus, for angles 40° and more 

above and below the horizontal plane, the position light intensity need be only 5 

percent of the light intensity in the horizontal plane. This provision thus recognizes 

that the significance of a position light decreases as zenith is approached. 

A provision similar to that being here established for the rear position light already 

exists with respect to the anticollision light. In this connection, minor visibility 

obstructions permitted in the rearward direction in the field of coverage of the 

anticollision light have been determined not to be detrimental to safety. 

286
04/12/05 AC 23-17B 

In light of the foregoing, obstructions within the dihedral angle in which the rear 

position light must show, which do not exceed 0.04 steradians in coverage and which 

occur within 30- of a vertical line through the rear position light, would not adversely 

affect safety. In addition, these amendments permitting minor obstructions in the 

field of coverage of rear position lights are consistent with the provisions of Secs. 

23.1385(c), 25.1385(c), 27.1385(c), and 29.1385(c) which recognize practicable 

considerations in the location of rear position lights.” 

Amendment 23-43 and Subsequent 

This amendment removed the words “forward and rear” from paragraph (a). These 

changes are necessary for compatibility with revised § 23.1385. 

287
AC 23-17B 04/12/05 

23.1389 Position light distribution and intensities 

The corresponding rule in CAR 3 is CAR 3.702. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.27. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “Section 

23.1389 is amended in paragraph (b) by removing words "Forward and rear" from 

the heading, by changing the word "position" in the heading to read "Position", and 

by removing the words "forward and rear" from the first sentence; in paragraph 

(b)(3) by removing the word "forward" in the last sentence and inserting in its place 

the words "left and right". 

These changes are necessary for compatibility with revised § 23.1385. 

288
04/12/05 AC 23-17B 

23.1391 Minimum intensities in the horizontal plane of position lights

The corresponding rule in CAR 3 is CAR 3.702. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.28. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “Section 

23.1391 is amended in the section heading by removing the words "forward and 

rear" and in the table by removing the words "(forward red and green)" and inserting 

in their place "(red and green)".” 

These changes are necessary for compatibility with revised § 23.1385. 

289
AC 23-17B 04/12/05 

23.1393 Minimum intensities in any vertical plane of position lights

The corresponding rule in CAR 3 is CAR 3.702. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.29. 

Amendment 23-43 and Subsequent 

This amendment removed the words “forward and rear” in the section heading. 

These changes are necessary for compatibility with revised § 23.1385. 

290
04/12/05 AC 23-17B 

23.1395 Maximum intensities in overlapping beams of position lights

The corresponding rule in CAR 3 is CAR 3.702. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.30. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “Section 

23.1395 is amended in the section heading by removing the words "forward and 

rear".”

This change is necessary for compatibility with revised § 23.1385. 

291
AC 23-17B 04/12/05 

23.1397 Color specifications 

The corresponding rule in CAR 3 is CAR 3.703. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.31.

Original Issue and Subsequent

See AC 20-74, “Aircraft Position and Anticollision Light Measurements,” for 

guidance on color measurements. 

Amendment 23-11 and Subsequent 

A proposed revision to NPRM 70-21 changed the color requirements for Aviation 

White as follows: “It is proposed to expand the chromaticity-coordinate range for 

the color Aviation White in order to provide for the use of white condenser-discharge 

anticollision lights, including Xenon types, in the implementation of proposal 1.”

292
04/12/05 AC 23-17B 

23.1399 Riding light 

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.704. 

There is no corresponding rule in the Airship Design Criteria. 

293
AC 23-17B 04/12/05 

23.1401 Anticollision light system 

The corresponding rule in CAR 3 is CAR 3.705. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.32.

Original Issue and Subsequent

If certification for night operation is sought, an anticollision light system, according 

to 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

A proposed revision to NPRM 70-21 explains this amendment as follows: “(1) It is 

proposed to amend Sec. 23.1401 (and the corresponding sections of Parts 25, 27, and 

29) to permit the use of either aviation red or Aviation White anticollision lights. The 

regulations now require that each anticollision light must be aviation red. However, 

neither the research conducted in the past by the FAA, and by others, nor the 

arguments submitted in response to Notice 70-7 by proponents of each color have 

conclusively established that one color is superior to the other at comparable 

intensities. Current FAA/NASA research may shed new light on this question. 

It is clear, however, that if the red color for anticollision lights were changed to 

white, the anticollision light intensity could be increased by a factor of 3 to 5 (without 

increasing the electric power consumption) by merely removing the red filter needed 

to meet current standards, or by replacing it with a white light (condenser discharge 

or other type) that uses about the same amount of power. Alternatively, if the 

regulation permitted Aviation White anticollision lights, the currently prescribed 

intensities could be produced with one-fifth to one-third of the electric power now 

being used by the red anticollision lights. Although it is true that the use of Aviation 

White could introduce a backscatter problem on some aircraft, we believe that this 

problem can be solved by the relocation of lights, by appropriate masking, or by 

other methods, as has already been done with respect to the white "supplementary" 

high-intensity lights installed on many aircraft now in service. 

294
04/12/05 AC 23-17B 

Some of the comments suggested that safety might be compromised by permitting 

aircraft using differently colored anticollision lights to operate in the same airspace. 

The FAA does not agree. Many aircraft currently operating in the United States do 

not now display red anticollision lights exclusively, since the current regulations 

permit the display of high-intensity flashing white lights (and even red and green 

lights) as "supplementary" lights. Frequently, the "supplementary" white lights are 

seen first as the aircraft is approached. The use of such "supplementary" lights has 

been permitted for more than 10 years and there has been no adverse effect on safety. 

Therefore, there is no reason to believe that the optional use of either red or white 

anticollision lights in the future would compromise safety in any way. 

On the other hand, since the standards of the International Civil Aviation 

Organization (ICAO) now specify red as the color for anticollision lights, the 

adoption of this proposal would make it necessary to file a formal notice of difference 

with ICAO. 

(2) It is proposed to expand the chromaticity-coordinate range for the color Aviation 

White in order to provide for the use of white condenser-discharge anticollision 

lights, including Xenon types, in the implementation of proposal 1. 

(3) It is proposed to increase the minimum effective intensities for anticollision lights 

and to make that increase applicable to both red and white lights on aircraft TC 

in the future. A majority of the persons responding to Notice 70-7 indicated that 

the currently prescribed intensity level for anticollision lights should be raised. 

In view of the fact that this increase would apply only to anticollision lights 

installed on aircraft for which an application for TC is made after the effective 

date of any final amendment containing this proposal, and in view of the current 

state-of-the-art in anticollision lights, the FAA considers that the application of 

the proposal of all aircraft is feasible. 

On the question of what increase in anticollision light intensity should be made, those 

who responded to Notice 70-7 offered suggestions that varied widely; but most 

recommended a four-fold increase over current levels. This corresponds roughly to 

the increase in intensity that would be attained by removing the red filter from 

existing anticollision lights, and is within the performance capability of state-of-theart condenser-discharge lights. The FAA believes this intensity recommendations has 

merit, and it is proposed to increase currently prescribed anticollision light 

intensities by a factor of four. 

295
AC 23-17B 04/12/05 

Contrary to the suggestion contained in various comments on Notice 70-7, the FAA 

believes that an intensity level standard that includes a specified infrared signal 

content for use with Pilot Warning Indicators (PWI) would be premature. The FAA 

considers that mandatory action on infrared signal content should await completion 

of current evaluations of the PWI system concept on civil aircraft. However, the 

proposed standard would not prevent any manufacturer from providing an infrared 

signal in his anticollision lights. 

There were also suggestions that the current anticollision light intensity should be 

retained for ground operation. However, the FAA does not consider that this is 

necessary since, under current rules, the anticollision lights may be dimmed or even 

turned off during ground operations.” 

This amendment is clarified by Final Rule, Docket 10129, as follows: “One 

commentator indicated that the FAA may have underestimated the problems 

associated with removal of the red filter on the anticollision light on existing aircraft. 

The commentator pointed out that it may be impossible to meet the current field-ofcoverage requirements because of the need to mask the unfiltered light to eliminate 

back-scatter. While the FAA recognizes this problem, it should be noted that it was 

not proposed to require anyone to remove the red filters on existing aircraft. 

Moreover, if an operator elected to do so, he would not be required to meet the new 

intensity requirements. Furthermore, he could relocate the light to minimize the 

back-scatter problem. Another comment expressed an objection to the proposed 

increase in the intensity level of anticollision light systems for future aircraft on the 

grounds that for small aircraft using red anticollision lights, power requirements 

would be unreasonable, service life short and reliability low, and that for small 

aircraft using the white anticollision lights, it would be possible to shield them for 

purposes of back-scatter without a reduction in the required field-of-coverage. The 

FAA is aware that for red anticollision lights more electrical power would be needed 

to meet the new requirements than has been provided in the past. However, the FAA 

believes that this additional power capacity can be provided on future aircraft at 

reasonable cost, without incurring a low-service-life or low-reliability penalty. 

Moreover, a manufacturer would now have the option of installing a white 

anticollision light, thereby eliminating the power problem. The back-scatter 

problems referred to by the commentator can be solved without diminishing the fieldof-coverage by installing a system consisting of three lights, one at each wing-tip and 

one on the tail.” 

There is no restriction on mixing aviation red and Aviation White anticollision lights 

on the same airplane. Also, there is no restriction on the ratio of red to white 

provided the light displayed in any one direction is either aviation red or Aviation 

White. 

296
04/12/05 AC 23-17B 

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 so 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 Whites “X” coordinate be no less 

than 0.300 and no greater than 0.540 (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) the measurement 

accuracy of the “X” value of chromaticity coordinates includes an error tolerance of 

plus or minus 0.008. It was not envisaged 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: 1) the 0.300 will not be exceeded more 

than 68 percent of the time and 2) 0.292 will not be exceeded 99.7 percent of the time 

(3 sigma), which includes the measurement error tolerance suggested by NBS. 

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this amendment as follows: “The 

present anticollision light rule in Sec. 23.1401(b) requires that the field of coverage 

extend in each direction within at least 30° above and 30° below the horizontal plane 

of the airplane. The FAA believes that this minimum coverage should be increased. 

Under the present rule, visibility of the light may be less than is needed when the 

airplane is approached by another aircraft ascending or descending, or when the 

airplane is banked at more than 30°. This proposal would expand the present field of 

coverage so that it extends in each direction within at least 75° above and 75° below 

the horizontal plane of the airplane. In addition, this proposal would specify at least 

a 20-candle effective intensity in the 30° to 75° field of coverage, thereby increasing 

the probability of seeing the airplane from other aircraft (including air rescue 

aircraft) and from control towers.” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 revises § 23.1401 to require installing an 

anticollision light system on all part 23 airplanes. Current § 23.1401 requires an 

anticollision light system only if certification for night operations is requested. 

297
AC 23-17B 04/12/05 

SAFETY EQUIPMENT 

23.1411 General 

The corresponding rules in CAR 3 are CAR 3.711, 3.713, 3.714 and 3.715. 

The corresponding rules in the Airship Design Criteria, FAA-P-8110-2, Change 2, are 

sections 6.33, 6.34 and 6.35. 

Amendment 23-17 and Subsequent 

A proposed revision to NPRM 75-10 explains this amendment as follows: “By 

amending Sec. 23.1411 in a manner substantively identical to that proposed for Sec. 

27.1411. The proposal would add a requirement for stowage provisions for all safety 

equipment in Sec. 27.1411 and would provide standards applicable to such stowage 

provisions.” 

Amendment 23-36 and Subsequent

There is no proposed change to § 23.1411 by the NPRM, but Final Rule, Docket 

25147, explains this amendment as follows: “One commenter suggests that the 

regulatory references in the notice should be crosschecked with the latest revisions of 

Part 23. As a result of reviewing the FAR, the FAA identified several sections of 

Part 23, which need to be changed for compatibility with the new requirements for 

dynamic testing of seat/restraint systems. The regulatory changes identified by this 

review are not substantive but provide consistency with the other changes made by 

this amendment. This review resulted in several additions to this final rule. 

Section 23.1411 is revised so that paragraph (b)(2) refers to section 23.561(b)(3).” 

298
04/12/05 AC 23-17B 

23.1413 Safety Belts and Harnesses [Removed] 

See guidance for section 23.785 in this AC. 

299
AC 23-17B 04/12/05 

23.1415 Ditching equipment

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.716, 3.717 

and 3.718. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.36. 

300
04/12/05 AC 23-17B 

23.1416 Pneumatic de-icer boot system 

The corresponding rule in CAR 3 is CAR 3.712. 

There is no corresponding rule in the Airship Design Criteria. 

Policy is available in AC 23.1419-2B. 

301
AC 23-17B 04/12/05 

23.1419 Ice protection 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.37. 

The § 23.1419 guidance in this AC is focused on the ice protection systems at the 

component and system level. For information about airplane performance and 

handling in icing conditions please refer to AC 23.1419-2B, “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 CAR 3 airplanes and 14 CFR, part 23 airplanes with a certification basis 

preceding Amendment 23-14, are allowed to fly in known icing conditions if their 

TCs do not include a prohibition against this operation. Some of these airplanes may 

be placarded against flight into known icing because they lack de-ice/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 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. This may require redundancy in ice protection system components to 

minimize hazards to the airplane if a probable malfunction or failure occurs. 

There is no requirement or allowance for making adjustments in the icing 

certification program for the frequency of encountering icing conditions. A 

likelihood 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.

302
04/12/05 AC 23-17B 

Amendment 23-43 and Subsequent

See ACs 23.1419-2, 23-143-1, 23-8 and 20-73. 

EASA AMC 23.1419 accepts FAA AC 23.1419-2 as an AMC to CS 23.1419. 

303
AC 23-17B 04/12/05 

MISCELLANEOUS EQUIPMENT 

23.1431 Electronic equipment 

The corresponding rule in CAR 3 is CAR 3.721. 

There is no corresponding rule in the Airship Design Criteria.

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 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 because of interaction of assignable 

frequencies in the NAS is not acceptable. 

Guidance for TCAS II installations is given in AC 20-131A, “Airworthiness 

Approval of Traffic Alert and Collision Avoidance Systems (TCAS II) and Mode S 

Transponders.” 

Automatic Air Navigation Facility (NAVAID) selection tuning (Auto-tune) of VHF 

Omnirange Station/Distance Measuring Equipment (VOR/DME) for flight 

management or multisensor navigation systems is designed to improve 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 from that wanted by the flight crew is a possibility. Visual cues 

indicating the auto-tune is still active may be 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

304
04/12/05 AC 23-17B 

installations that 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 using 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 and LORAN-C) 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”. It should 

be 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. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal would include electronic equipment that is being installed in part 23 

airplanes as well as radio equipment. When the existing regulation was adopted, 

radio equipment was the primary electronic equipment installed. For 

standardization in the application of FAA requirements, this proposal is consistent 

with Section 25.1431(a) and (c). Section 23.1309(b)(1) and (2) that are referenced 

are the proposed regulations in Notice 5, Small Airplane Airworthiness Review 

Program.” 

This amendment is clarified by Final Rule, Docket 26344, as follows: “This 

proposes to amend Sec. 23.1431 to revise the current rule that addresses radio 

equipment only by including other electronic equipment that is installed in a part 23 

airplane. Two comments were received. One commenter asks for a definition of the 

words, "critical environmental conditions" used in proposed Sec. 23.1431(a). 

Critical environmental conditions are those environmental conditions under which a 

piece of equipment will not perform its intended function. By including this 

requirement, conditions that may be critical to the operation of a piece of equipment 

must be considered. Consideration of such conditions would include, but not be 

limited to, temperature extremes, vibration levels, and humidity. 

The other commenter agrees with the proposal and suggests that Sec. 23.1431 be 

expanded to cover communications between pilots, radio transmission switches, and 

the effectiveness of aural warnings when headsets are being worn. Because these 

suggested expansion items were not included in the notice, their addition would be 

beyond the scope of the NPRM. This proposal is adopted as proposed.”

305
AC 23-17B 04/12/05 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this change to Amendment 23-49 as 

follows: “This proposal would add three new paragraphs to Sec. 23.1431. Proposed 

new paragraph (c) would require that airplanes required to be operated by more than 

one flightcrew member must be evaluated to determine if the flightcrew members can 

converse without difficulty when they are seated at their duty stations. Accident 

investigations have shown that, in some instances, conversation between the 

flightcrew members was severely hindered by the noise level in the cockpit and that 

the inability to communicate contributed to the accident. If the required evaluation 

shows that the noise level does not impair conversation, no further action is required. 

However, if the evaluation shows that conversation will be difficult, an 

intercommunication system would be required. 

Proposed new paragraph (d) would require that if installed communication 

equipment includes any means of switching from receive to transmit, the equipment 

must use "off-on" transmitter switching that will ensure that the transmitter is turned 

off when it is not being used. Transmitting equipment that remains in the transmit 

mode when not being used blocks the frequency being used and can create an unsafe 

condition by preventing other needed communication. 

Proposed new paragraph (e) would require that if provisions for the use of 

communications headsets are provided, it must be demonstrated that flightcrew 

members can hear aural warnings when a headset is being used. Aural warnings are 

required to warn the pilot of a condition that necessitates the pilots taking action; 

therefore, it is necessary to ensure that such warnings would be effective even when 

headsets are being used. 

During the development of the proposed new requirements in paragraphs (c) and (e),

the FAA considered proposing a requirement that compliance demonstrations should 

be conducted under actual cockpit noise conditions when the airplane is being 

operated. The FAA, however, ultimately determined that such a requirement could 

result in demonstrations conducted under more severe noise conditions than needed. 

Accordingly, no such requirement is being proposed. If the FAA determines in the 

future that noise conditions for demonstrations need to be specified, the FAA will 

define these conditions in advisory material.” 

The proposal to include testing under actual noise conditions in paragraphs (c) and (e) 

was revised by Final Rule, Docket 27806, as follows: “This proposal would add 

three new paragraphs to Sec. 23.1431. Proposed new paragraph (c) would provide 

that airplanes required to be operated by more than one flightcrew member be 

evaluated to determine if the flightcrew members can converse without difficulty 

when they are seated at their duty stations. Proposed new paragraph (d) would 

require installed communication equipment to use "off-on" transmitter switching that 

will ensure that the transmitter is turned off when it is not being used. Proposed new

306
04/12/05 AC 23-17B 

paragraph (e) would require that, if provisions for communication headsets are 

provided, the applicant must demonstrate that flightcrew members will receive all 

warnings when a headset is being used. The demonstration must be made under 

actual cockpit noise conditions. To clarify the conditions under which these 

evaluations should be conducted, not withstanding earlier harmonization agreements, 

these two paragraphs are being revised to include the phrase, "under actual cockpit 

noise conditions when the airplane is being operated." 

EASA AMC 23.1431(e) is acceptable for FAA certification. 

307
AC 23-17B 04/12/05 

23.1435 Hydraulic systems

The corresponding rule in CAR 3 is CAR 3.3.726. 

The corresponding rules in the Airship Design Criteria, FAA-P-8110-2, Change 2, are 

sections 6.4 and 6.38. 

Amendment 23-7 and Subsequent 

A proposed revision to NPRM 67-14 revised paragraph (c) to add the words "or 

propeller" after the words "part of an engine.” 

Amendment 23-14 and Subsequent 

A proposed revision to NPRM 71-13 explains this amendment as follows: “This 

proposal is appropriate because of increased reliance on hydraulic systems and the 

increasing complexity or such systems in Part 23 airplanes.” 

The NPRM was revised by Final Rule, Docket 11011, as follows: “Sec. 

23.1435(a)(2) has been revised to indicate that the required indication must be to the 

flight crew.” 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: 

“Conference proposal 469 recommended adding to Section 23.1435(c) a requirement 

that propeller unfeathering accumulators be considered as an integral part of the 

propeller and small (1 quart max.) nonpressurized reservoirs be acceptable. 

Propeller unfeathering accumulators have been accepted as an integral part of a 

propeller. The conference discussion supported clarification of the requirement and 

the allowance of some small accumulators, such as for the brake systems on singleengine airplanes. The FAA has further considered these issues and concludes such 

accumulators should be allowed provided their total capacity is limited to one quart 

or less.”

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 clarifies paragraph (c) with the following 

explanation: “Since the close of the comment period for the Small Airplane 

Airworthiness Review Program Notice No. 3 (55 FR 40598, October 3, 1990), now 

adopted by Amendment No. 23-43 (58 FR 18958, April 9, 1993), the FAA has been 

involved in discussions of the installation of hydraulic accumulators that are 

permitted by Sec. 23.1435(c). These discussions have shown that applicants are 

likely to find Sec. 23.1435(c) difficult to understand because of the way it is worded. 

308
04/12/05 AC 23-17B 

This notice would further revise Sec. 23.1435(c) to clarify under what circumstances 

a hydraulic accumulator and reservoir may be installed on the engine side of any 

firewall.”

309
AC 23-17B 04/12/05 

23.1437 Accessories for multiengine airplanes

No policy available as of September 30, 2003. 

This rule was adopted on February 1, 1965 as a recodification of CAR 3.725. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.39. 

310
04/12/05 AC 23-17B 

23.1438 Pressurization and pneumatic systems

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.40.

Amendment 23-20 and Subsequent 

A proposed revision to NPRM 75-23 explains this rule as follows: “Components 

(such as ducts and couplings) of pressurization and pneumatic systems have failed at 

an unacceptable rate in service. The proposed standards for these components have 

been effective in preventing design deficiencies in the past.” 

The rule proposed in NPRM 75-23 was revised by Final Rule, Docket 14625, and is 

explained as follows: “For comments related to the proposal to add a new Section 

23.1438, and for an explanation of revisions made to proposed Section 23.1438, see 

Proposal 5-33. Proposal 5-33 includes: One commentator stated that if the term 

"components" in proposed Secs. 25.1438(a) and (b) included all parts of the system it 

would not be compatible with other Secs. of Part 25, which use the term "elements". 

The FAA intended that the term "components" include all parts of the system. 

Therefore, in order to be consistent with current Section 25.1435, which uses the term 

"elements", that term is substituted for the term "components" in proposed 

Secs. 25.1438(a) and (b).

One commentator questioned the need to specify a higher burst pressure for 

pneumatic systems than for pressurization systems, contending that pneumatic 

systems do not necessarily operate at higher pressure and that some are derived from 

pressurization systems. The FAAs experience has been that pneumatic systems in 

airplanes are operated at higher pressures, even when a common pressure source is 

provided for both pneumatic and pressurization systems. 

One commentator objected to the provision in proposed Section 25.1438(b) requiring 

a burst pressure test of 4.0 times maximum normal operating pressure, contending 

that the industry has historically designed and tested pneumatic systems to a burst 

pressure of 3.0 times maximum normal operating pressure and that service 

experience over millions of flight hours has proven the integrity of those systems. 

Another commentator pointed out that pneumatic deicer boots in general use today 

cannot sustain a pressure of 4.0 times maximum normal operating pressure. The 

FAA agrees with these comments, and proposed Section 25.1438(b) is revised to 

specify a burst pressure of 3.0 times maximum normal operating pressure. 

One commentator suggested revision of proposed Section 25.1438 to allow the use of 

analysis, or a combination of analysis and test, as an alternative method of 

compliance to eliminate unnecessary testing. The FAA agrees that there are 

311
AC 23-17B 04/12/05 

instances where an analysis, or a combination of analysis and test, may be equivalent 

to a test under proposed Section 25.1438(a) or (b). Accordingly, a new Section 

25.1438(c) is added to provide this alternative.” 

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 enough overstrength material elements are 

part of a design, then reduced material strength because of 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 show structural property phenomenon and structural strength 

changes because of temperature changes. Certain composite materials may be 

characteristic of adverse performance in certain temperature environments. Also, 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. 

312
04/12/05 AC 23-17B 

23.1441 Oxygen equipment and supply

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Amendment 23-9 and Subsequent

A revision to NPRM 67-30 proposed this rule based on the following explanation: 

“In a study concerning the physiological requirements for oxygen entitled "Oxygen in 

General Aviation" dated December 1964, the Office of Aviation Medicine of the FAA 

reported that, while individual oxygen requirements may be influenced by factors 

such as fatigue, age, use of drugs and tobacco, night, etc., nevertheless supplemental 

oxygen, dependent on altitude, is a physiological necessity for all humans. For this 

reason and because of the increasing number of small airplanes having high altitude 

capability, there appears to be a need for airplane airworthiness standards and for 

general operating rules relating to oxygen system requirements. 

In the past, type certification of small airplanes with oxygen equipment installed has 

been handled on a case-by-case basis either by application of special conditions or 

under the general operating limitations of Sec. 23.1501. This amendment proposes 

uniform airworthiness standards for oxygen equipment and supply. It also proposes 

performance standards relating to oxygen mass flow and equipment standards for 

oxygen dispensing units. 

Parts 121 and 135 of the Federal Aviation Regulations currently prescribed oxygen 

requirements for pilots and passengers. For consistency in the interest of safety, all 

other aircraft operating in the same airspace and environmental conditions should be 

under comparable standards. The FAA, therefore, considers it appropriate to 

propose general operating and flight rules governing the use of supplemental oxygen 

in pressurized and unpressurized cabin aircraft. 

FAA studies have indicated that for any person in a pressurized cabin subjected to 

decompression at the higher altitudes, the short time available to don an oxygen mask 

before losing consciousness may be critical. In such case, the smaller the cabin 

volume, the more aggravated the situation becomes. As an added operational 

consideration, the FAA, therefore, believes it necessary for safety, at flight altitudes 

above 35,000 feet, that at least one pilot at the controls be required to wear an 

oxygen mask. The corresponding altitude for Part 121 operations had originally 

been established at 35,000 feet. However, it was subsequently raised, in steps, to 

41,000 feet, on the basis of the extensive operating experience with the large 

transport-category airplanes. This satisfactory operating experience is attributable, 

in part, to the stringent Part 121 crew training and maintenance requirements. Such 

compensating conditions do not presently exist for operations conducted under 

Part 91 and the proposed cutoff altitude has been accordingly set at 35,000 feet.

313
AC 23-17B 04/12/05 

A new Sec. 91.32 would be added to the general operating rules of Part 91 to 

regulate the use of supplemental oxygen throughout the range of cabin altitudes and 

specify the oxygen supply to be carried, including the oxygen necessary for 

emergency descent. Supplemental oxygen refers to any oxygen, by whatever means 

supplied, that is furnished in addition to the oxygen normally present in the air.” 

Amendment 23-9 is clarified by Final Rule, Docket 8281, as follows: “Several 

commentators objected to the proposal on the basis that it would require the owners 

of small airplanes that are not designed to operate at altitudes where oxygen is 

required to purchase expensive oxygen equipment for their airplanes. This, however 

is not the case. The proposal would not require the installation of oxygen equipment 

on all small airplanes. Oxygen equipment would have to be installed only on small 

airplanes that are to be used in operations conducted at the altitudes for which 

supplemental oxygen must be provided and used in accordance with the requirements 

of Part 91. The need for oxygen in such operations far outweighs any economic 

burden that may be involved in the installation and maintenance of the oxygen 

equipment. 

Several commentators suggested revising proposed Sec. 23.1441(a) which states that 

there must be a means to allow the crew to readily determine during flight the 

quantity of oxygen available in each source of supply. It was recommended that the 

proposal be changed to specify a means for determining the remaining number of 

hours of available oxygen. These commentators are primarily concerned with the 

pressure gauges being furnished with oxygen systems. They believe that a pressure 

gauge is not an adequate instrument for determining the quantity of oxygen available. 

The FAA does not agree. In most instances, a pressure gauge in the oxygen system 

will indicate the quantity of oxygen available in the source of supply. If, in any 

particular installation, it is determined that a pressure gauge would not satisfy the 

requirement, then some other means to allow the crew to determine, during flight, the 

quantity of oxygen available in each source of supply would have to be provided. In 

any event, the remaining number of hours of available oxygen can be readily 

obtained once the quantity of oxygen available in each source of supply is known. 

In response to numerous comments, Sec. 23.1441 has been revised to make it clear 

that portable oxygen equipment may be used to meet the supplemental oxygen 

requirements.” 

Plastic lines (nylon, 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: 

314
04/12/05 AC 23-17B 

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,” and American Society for Testing and Materials (ASTM) 

Manual 36, “Safe Use of Oxygen and Oxygen Systems, Guidelines for Oxygen 

System Design, Materials Selection, Operations, Storage, and Transportation”, 

dated 2000Chapter 6; for additional guidance material. 

Design and install oxygen tubes/hoses in such a manner 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) the installed 

tubes/hoses will not vibrate/flap in a manner that would either defeat the chaffing 

prevention wrap, or to impact other system elements such as sensors, wires, and 

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 need 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 as follows: (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. 

If an airplane is delivered without an oxygen system, its AFM should have a 

limitation or there should be a placard prohibiting flight above 14,000 feet 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. 

315
AC 23-17B 04/12/05 

The rule allows portable systems to be designed by a TC holder subject to actions to 

prevent the portable system from being a hazard by not being properly secured in an 

emergency landing (according to § 23.561). The precise requirement for portable 

systems should be determined for the specifics of the airplane type (that is, its service 

ceiling and number of seats). The rulemaking history does not state a safety 

deficiency in current methods of providing supplemental oxygen by operatorprovided, portable, constant flow oxygen systems either attached to the seatback or 

restrained by a seatbelt. 

If a TC 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 according to § 23.561 either by attachment to a 

seatback or a seatbelt. This could be commercially available or produced by the TC 

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. 

Amendment 23-43 added § 23.1441(e). It requires there be a means readily available 

to the flight crew to both turn ON and turn OFF the oxygen supply at the highpressure source. A flow fuse that senses a rupture in an oxygen line and 

automatically shuts off is not an ELOS to the required pilot operated switch since it 

only applies to one failure where oxygen shutoff is needed. It was made clear by the 

NPRM the switch was needed to turn ON and OFF the oxygen system, to shutoff if a 

rupture or fire occurs. 

316
04/12/05 AC 23-17B 

23.1443 Minimum mass flow of supplemental oxygen

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Original Issue and Subsequent

When there is full compliance to this regulation, there is no need to consider the 

likelihood of a pressurization failure or to require an immediate descent in altitude if 

a failure occurs. 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 ELOS finding. 

Amendment 23-9 and Subsequent 

A proposed revision to NPRM 67-30 explains this amendment as follows: “In 

connection with the standards for minimum mass flow of supplemental oxygen, the 

proposed Sec. 23.1443 contains requirements for continuous flow but not for demand 

flow oxygen equipment. The requirements of the more commonly used continuous 

flow systems are thus covered, while the mass flow of demand systems, where such 

systems are used, would be approved on a case-by-case basis as before.” 

This amendment is clarified by Final Rule, Docket 8281, as follows: “One 

commentator suggested incorporating a provision in Sec. 23.1443 which would 

require compliance with the proposed graph (depicting oxygen mass flow rates) only 

up to the cruising altitude for which type certification of the airplane is desired. The 

recommendation is premised on the contention that it is unnecessary and 

economically unreasonable to install oxygen equipment capable of a specified mass 

flow rate at 40,000 feet if the airplane will be operated at a much lower altitude. The 

purpose of Sec. 23.1443 is to show the minimum oxygen mass flow rates which would 

be required for each occupant of an airplane at the corresponding cabin pressure 

altitudes. The applicant for a type certificate would not be required to install oxygen 

equipment capable of supplying the minimum oxygen mass flow rates for the entire 

range of cabin pressure altitudes shown in the graph unless his airplane is capable of 

operating through the entire range. Accordingly, Sec. 23.1443 is revised to state that 

the oxygen equipment installed in the airplane must be capable of supplying to each 

occupant the appropriate flow of oxygen for all altitudes up to and including the 

maximum operating altitude of the airplane.” 

317
AC 23-17B 04/12/05 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal modifies the oxygen flow rates for small airplanes. These new requirements 

are needed for current and future airplanes that will be certificated to higher 

altitudes where oxygen is required. 

Conference proposal 472 recommended revising Section 23.1443 by designating the 

present paragraph as paragraph (a) and adding a new paragraph (b) to read 

essentially the same as Section 25.1443(b). The justification given was that "The 

oxygen flow rate requirements of parts 23 and 25 are different. However, both 

regulations provide requirements needed to ensure continuous flow rates up to cabin 

pressure altitudes of 40,000 feet. Proposed Section 23.1443 is a combination of these 

requirements. Proposed Section 23.1443(a) contains current Section 23.1443 with 

regard to continuous flow requirements, and Section 23.1443(b) is derived from 

Section 25.1443 for demand system requirements. The language of the recommended 

change would allow Section 23.1443 to cover both continuous flow oxygen and 

demand systems without interpretation from part 25." 

“Conference proposal 473 recommended revising § 23.1443 essentially as shown in 

this proposal since the oxygen flow rate requirements of parts 23 and 25 are 

different. However, both regulations provide requirements needed to assure 

continuous flow rates up to cabin pressure altitudes of 40,000 feet. Proposed 

§ 23.1443 is a combination of these requirements. Sections 23.1443 (a)(1) and (a)(2)

contain the continuous flow requirements of part 25 and allow the applicant to 

comply with those requirements or with paragraph (a)(3), which is the current 

continuous flow requirement of part 23. Demonstrating compliance with proposed 

§ 23.1443(a)(3) is easier, but results in a larger volume of oxygen and more weight. 

Demonstrating compliance with § 23.1443 (a)(1) and (a)(2) is harder, but results in a 

lesser volume of oxygen. By allowing the applicant to choose either method of 

compliance, this requirement permits freedom of design. 

When presented for comment at the conference, the FAA confirmed that this 

conference proposal would allow alternatives of continuous flow oxygen equipment 

or demand oxygen equipment. With a good face-fitting mask, less oxygen will be 

used with a demand system than with a continuous flow system. Studies on altitude 

sickness and the impairment of ability to function on continuous flow equipment at 

altitudes above 25,000 feet leads the FAA to reconsider this issue, especially for 

flight crews. 

Proposed paragraph (c) provides the flow rate requirements for first-aid oxygen 

equipment if installed, but does not require its installation. These requirements are 

identical to the first-aid oxygen flow rate requirements in part 25. With the recent 

addition on commuter category airplanes in part 23, first-aid oxygen equipment is 

more likely to be installed in part 23 airplanes.

318
04/12/05 AC 23-17B 

Proposed paragraph (d) is clarifying by providing definitions of the term "BTPS" and 

"STPD" as used in this section. 

Post conference review of these comments and the oxygen requirements of parts 91, 

121, and 135 led to the conclusion that (1) adding the equivalent of the part 25 

oxygen requirements to part 23 will provide adequate protection for both flight crew 

and passengers; and (2) that crewmembers should have demand oxygen equipment 

for operations above 25,000 feet.” 

This amendment is clarified by Final Rule, Docket 26344, as follows: “This 

proposes to amend Sec. 23.1443 to modify the oxygen flow rates for part 23 airplanes 

by providing alternate procedures that may be used to substantiate satisfactory 

continuous flow oxygen equipment. One commenter requested that the FAA make it 

clear that the 40,000 foot altitude limit in this proposal is not an absolute altitude 

limit for part 23 airplanes. The FAA agrees. The altitude limit in this proposal does 

not constitute an absolute altitude limit for the approval of part 23 airplanes; 

however, the approval of individual airplanes would be limited to those altitudes 

where safe occupant protection is provided.” 

319
AC 23-17B 04/12/05 

23.1445 Oxygen distribution system 

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Amendment 23-43 and Subsequent

A proposed revision to NPRM 90-23 explains this amendment as follows: “This 

proposal will establish standards for oxygen distribution systems not heretofore 

required. These requirements will prevent installation of plastic hoses in pressurized 

oxygen systems. 

Conference proposal 474 recommended adopting equipment standards for oxygen 

systems essentially the same as proposed here. The justification given was that 

several accidents have occurred in airplanes where nylon tubing was used in an 

oxygen system pressurized to 70 psi. Because oxygen can support vigorous 

combustion, oxygen system installations warrant special attention in certification 

programs. 

When presented for comment at the conference, the two commenters agreed that 

pressurized plastic tubing is inappropriate for oxygen system but did not completely 

agree with the proposal because they believed the rules should not specifically 

preclude all nonmetallic tubing. Some composite airplanes may need nonmetallic 

oxygen lines for lightning strike protection. In view of these comments, a phrase that 

allows a showing of suitability to the installation was added to the proposal.” 

The guidance in this AC for § 23.1441, Amendment 23-9 and subsequent, for plastic 

lines is applicable to this regulation. 

320
04/12/05 AC 23-17B 

23.1447 Equipment standards for oxygen dispensing units 

There is no corresponding rule in CAR 3 or the Airship Design Criteria.

Amendment 23-9 and Subsequent 

This amendment is clarified by Final Rule, Docket 8281, as follows: “It was 

recommended that paragraphs (b) and (c) of proposed Sec. 23.1447 be revised to 

specify that they are applicable only to pressurized airplanes. The commentator 

asserts that these proposals should not apply to nonpressurized airplanes, which are 

not subject to sudden-decompression occurrences. The FAA agrees and this has been 

made clear in the final rule. In addition, further FAA studies have shown that for 

those pressurized aircraft designed to operate at 25,000 feet MSL or below, the pilot 

will have sufficient time to descend to safer altitudes prior to the onset of any 

symptoms of hypoxia in the event of a sudden decompression. For this reason, the 

proposal in paragraph (b) of Sec. 23.1447, requiring an oxygen supply terminal and 

unit of oxygen dispensing equipment within reach of each member of the required 

minimum flight crew for those airplanes designed to operate at flight altitudes up to 

and including 25,000 feet, is unnecessary and has not been adopted.” 

The Small Airplane Directorate was recently asked for applicable policy about 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) was incorporated into part 23 by 

Amendment 23-9 and item (2) above by 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 AC 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, either as upright, reclined, swiveled or tracked. The mask need 

not be presented in front of all people if there is sufficient “crowd awareness,” that is, 

the vast majority has proper presentation and the others can readily see the masks 

have been presented. These latter people 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 

seating positions result in many different group combinations. Each combination 

should have an adequate number of masks reachable by every person. Consideration 

321
AC 23-17B 04/12/05 

should be given to minimizing the likelihood of people 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. 

Either 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

A proposed revision to NPRM 75-23 explains this amendment as follows: “The 

proposal would require automatic presentation of oxygen dispensing units on 

airplanes certificated under Part 23 for operations above 30,000 feet. Should loss of 

pressurization occur at those altitudes, the time of useful consciousness varies 

between 90 seconds at 30,000 feet and 10-12 seconds at 40,000 feet, necessitating 

automatic presentation of oxygen dispensing units. It is also proposed to specify 

14,000 feet as the cabin pressure altitude at which they must be automatically 

presented so that they would be ready for use not only when the cabin pressure 

altitude goes very rapidly to very high values but also in the event of a pressurization 

system failure or malfunction after which (under present Sec. 23.841(a)) the cabin 

pressure altitude may go to 15,000 feet. The 14,000-foot cabin pressure altitude for 

automatic presentation is also being proposed for Part 25. 

In addition, since service experience has shown that the automatic presentation 

feature may fail, it is proposed that a manual means be provided to enable each 

occupant to gain access to his oxygen dispensing unit. This manual means is also 

being proposed for Part 25.” 

This amendment is clarified by Final Rule, Docket 14625, as follows: “For 

comments related to the proposal to add new Secs. 23.1447(c) and (d), and for an 

explanation of the revisions made to proposed Secs. 23.1447(c) and (d), see Proposal 

5-34. One commentator objected to the provision in proposed Section 25.1447(c)(1)

which would require that oxygen dispensing units be automatically presented before 

the cabin pressure altitude exceeds 14,000 feet, contending that long-standing FAA 

policy has been that the altitude for automatic presentation should be 15,000 feet and 

service experience over the last 16 years has not shown a need to reduce that 

altitude. The commentator further stated that the flight crew is given a warning when 

or before the cabin pressure altitude reaches 10,000 feet and is therefore alerted (in 

the event of a gradual increase in cabin pressure altitude) to the need for appropriate 

action either to maintain a safe cabin pressure altitude or manually deploy the 

dispensing units. Another commentator suggested that the presentation altitude be 

14,500 feet, rather than 14,000 feet, to take equipment tolerances into account. 

322
04/12/05 AC 23-17B 

In light of the comments received and after further review, the FAA believes that 

there is insufficient evidence at the present time to justify a requirement for the 

automatic presentation of oxygen dispensing units before the cabin pressure altitude 

exceeds 14,000 feet, and that the widely-used value of 15,000 feet provides an 

adequate level of safety. Proposed Section 25.1447(c)(1) is revised accordingly. 

Several commentators disagreed with the provision in proposed Section 

25.1447(c)(1) that would require that each occupant be provided with a manual 

means to make the oxygen dispensing unit immediately available, contending that 

manual back-up for the automatic presentation system should be provided for use by 

the crew only, to avoid tampering by the passengers. Another commentator stated 

that the average passenger would not be capable of operating such manual means 

properly and quickly, and that manual means are not feasible for ceiling or hatrack 

mounted dispensing units. 

The FAA believes that a manual means must be provided to back up the automatic 

presentation system, but is persuaded that it may not be in the interest of safety to 

require that a manual means be provided for passengers. Accordingly, proposed 

Section 25.1447(c)(1) is revised to require only that a manual means for the 

development of the dispensing units be provided for the crew.” 

Amendment 23-30 and Subsequent 

This amendment allows the use of nasal cannulas for operation up to an altitude of 

18,000 feet MSL. These are simple devices with no known service problems, and the 

FAA has not developed a design standard for them. 

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. 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. 

Amendment 23-43 and Subsequent 

A proposed revision to NPRM 90-23 explains this change to Amendment 23-43 as 

follows: “This proposal would add presentation requirements for the demand 

oxygen equipment required by Section 23.1441(d) and allow the option of quickdonning type oxygen dispensing units.”

323
AC 23-17B 04/12/05 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this amendment as follows: 

“Proposed new Sec. 23.1447(a)(4) would require that if radio equipment is installed 

in an airplane, flightcrew oxygen dispensing units must be designed to allow the use 

of communication equipment when oxygen is being used. If radio equipment is 

installed, that equipment cannot perform its intended function if the flightcrew is not 

provided the proper means for its utilization under all operating conditions, 

including operations when oxygen is being used. 

This proposal would not require all flightcrew oxygen dispensing units to be 

equipped with communication equipment. Since an airplane may be operated in 

uncontrolled airspace, where two-way radio communication is not required and, at 

the same time, be at altitudes where oxygen is required for the flightcrew members, 

some airplanes have a crew oxygen system but no radio equipment. It would be 

inappropriate to require the flightcrew dispensing units of those airplanes to be 

equipped with communication equipment. 

The proposed revisions to Sec. 23.1447(d) would require the flightcrew oxygen 

dispensing units to be automatically presented before the cabin pressure altitude 

exceeds 15,000 feet or be the quick-donning type if the airplane is certificated for 

operation above 25,000 feet. The requirement in paragraph (e) for the passenger 

dispensing units to be automatically presented if the airplane is approved for 

operation above 30,000 feet has not been revised. The revision to paragraph (d)

would provide the flightcrew and the airplane passengers the same level of safety as 

provided by other airworthiness standards. This proposed revision is also consistent 

with the proposed revision of Sec. 23.841 in this notice.” 

324
04/12/05 AC 23-17B 

23.1449 Means for determining use of oxygen

No policy available as of September 30, 2003. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

325
AC 23-17B 04/12/05 

23.1450 Chemical oxygen generators

No policy available as of September 30, 2003. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

326
04/12/05 AC 23-17B 

23.1451 Fire protection for oxygen equipment

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this rule as follows: “This proposed 

new section would specify that fire protection is needed for oxygen equipment 

installations. Section 23.1451(a) and (b) would, respectively, prohibit the installation 

of oxygen equipment in designated fire zones and require that oxygen system 

components be protected from the heat from designated fire zones. 

Proposed Sec. 23.1451(c) would require oxygen equipment and lines to be separated 

from other equipment or to be protected in a manner that would prevent escaping 

oxygen from striking grease, fluids, or vapors. The impingement of pure oxygen on 

certain materials will lower their combustion point to a value where ignition will 

occur in ambient conditions thereby creating a potential source for an airplane fire. 

In one instance, an airplane was destroyed by fire that resulted when escaping 

oxygen impinged on lubricating material during maintenance of the airplane. The 

proposed new section would ensure that oxygen systems are protected to prevent fire 

hazards that can result from escaping oxygen.” 

327
AC 23-17B 04/12/05 

23.1453 Protection of oxygen equipment from rupture

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 explains this rule as follows: “This proposed 

new section would clarify the rupture protection needed for oxygen system 

installation. Rupture protection for oxygen systems is currently required by the 

application of the structures load requirements of part 23. The addition of 

Sec. 23.1453(a) would clarify the application of these load requirements and would 

identify the need to consider maximum temperatures and pressures that may be 

present. Section 23.1453(b) would identify the protection to be provided for high 

pressure oxygen sources and the high pressure lines that connect such sources to the 

oxygen system shutoff valves.” 

Amendment 23-49 is clarified by Final Rule, Docket 27806, as follows: “Proposed 

new Sec. 23.1453 would clarify the rupture protection needed for oxygen system 

installation. Rupture protection for oxygen systems is currently required by the 

application of the structure load requirements of part 23. The addition of 

Sec. 23.1453(a) would clarify the application of these load requirements and would 

identify the need to consider maximum temperatures and pressures that may be 

present. Section 23.1453(b) would identify the protection to be provided for high 

pressure oxygen sources and the pressure lines that connect such sources to the 

oxygen system shutoff valves. 

The comments received on this proposal from the JAA and the Civil Aviation 

Authority (CAA). CAA noted that the word "high" in paragraph (b) could lead to 

confusion and require interpretation. Accordingly, they suggested that the words 

"High pressure oxygen sources" be revised to read as follows: "Oxygen pressure 

sources." This is the same text that is used in JAR 23. 

The FAA agrees with the suggested wording change. When the proposal was 

originally drafted, the FAA was considering the oxygen source side of the oxygen 

regulator, the high pressure side, and the passenger dispensing side of the regulator, 

the low pressure side; thus, the word "high" was used. 

The suggested change will not alter the requirements applicability and will be more 

clearly understood. It is also noted that the suggested text change will more closely 

align with the same requirement in Sec. 25.1453. Section 23.1453 is changed by 

revising the first four words of proposed paragraph (b) to read, "Oxygen pressure 

sources. 

This section is adopted with the above change.” 

328
04/12/05 AC 23-17B 

23.1457 Cockpit voice recorders

EASA AMC 23.1459(b) is acceptable for FAA certification. 

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-35 and Subsequent 

This rule is explained by Final Rule, Docket 25530, as follows: “Significant changes 

between the proposed rule and the final rule are summarized below. These changes 

were discussed in the appropriate sections of the Comment Discussion above: 

(1) The compliance dates have been extended from 2 years to 3 years. 

(2) Pre-wiring requirements in Sec. 91.35 and Sec. 135.52 have been deleted. 

(3) The requirements for airplanes currently equipped with a digital date bus are 

clarified in Secs. 121.343(e) and 125.225(c). 

(4) Certain of the parameters in Appendix B of Part 121 have been made optional. 

(5) Aircraft required to be upgraded under Sec. 135.152(b) must meet the 11-

parameter or 17-parameter flight recorder requirements (currently required of 

certain airplanes used under Part 121) except those manufactured after 

October 11, 1991. 

(6) Requirements pertaining to the use of boom microphones have been clarified. 

(7) The resolution readout column in each of the appendices has been made 

applicable only to aircraft manufactured after October 11, 1991. 

The changes above were coordinated with the NTSB staff who have indicated 

agreement that this final rule meets the intent of the NTSB recommendations in its 

letter of June 19, 1987. 

In addition the above changes, Parts 23 and 27 have been revised to include 

airworthiness requirements for flight and voice recorders. In the proposal such 

airworthiness requirements were only in Parts 25 and 29, and all operators subject to 

flight and voice recorder requirements were referred to those parts. In the interest of 

consistency, the final rule includes airworthiness requirements for flight and voice 

recorders in Parts 23, 25, 27, and 29. References in the operating rule have been 

revised accordingly.” 

329
AC 23-17B 04/12/05 

23.1459 Flight recorders

There is no corresponding rule in CAR 3 or the Airship Design Criteria. 

Amendment 23-35 and Subsequent 

Significant changes between the proposed rule and the final rule are summarized 

below. These changes were discussed in the appropriate sections of the comment 

discussion above: 

“(1) The compliance dates have been extended from 2 years to 3 years. 

(2) Pre-wiring requirements in Sec. 91.35 and Sec. 135.52 have been deleted. 

(3) The requirements for airplanes currently equipped with a digital date bus are 

clarified in Secs. 121.343(e) and 125.225(c). 

(4) Certain of the parameters in Appendix B of Part 121 have been made optional. 

(5) Aircraft required to be upgraded under Sec. 135.152(b) must meet the 11-

parameter or 17-parameter flight recorder requirements (currently required of 

certain airplanes used under Part 121) except those manufactured after October 

11, 1991. 

(6) Requirements pertaining to the use of boom microphones have been clarified. 

(7) The resolution readout column in each of the appendices has been made 

applicable only to aircraft manufactured after October 11, 1991. 

The changes above were coordinated with the NTSB staff that has indicated 

agreement that this final rule meets the intent of the NTSB recommendations in its 

letter of June 19, 1987. 

In addition the above changes, Parts 23 and 27 have been revised to include 

airworthiness requirements for flight and voice recorders. In the proposal such 

airworthiness requirements were only in Parts 25 and 29, and all operators subject to 

flight and voice recorder requirements were referred to those parts. In the interest of 

consistency, the final rule includes airworthiness requirements for flight and voice 

recorders in Parts 23, 25, 27, and 29. References in the operating rule have been 

revised accordingly.” 

330
04/12/05 AC 23-17B 

23.1461 Equipment containing high energy rotors 

There is no corresponding rule in CAR 3. 

The corresponding rule in the Airship Design Criteria, FAA-P-8110-2, Change 2, is 

section 6.41.

Amendment 23-20 and Subsequent

A proposed revision to NPRM 75-23 explains this rule as follows: “By adding a new 

Sec. 23.1461, following Sec. 23.1449, that would be substantively identical to the 

proposed new Sec. 25.1461. By adding a new Sec. 25.1461: This proposal would 

add requirements for protection against the failure of equipment containing high 

energy rotors, such as turbine engine starters, air cycle machines, and certain 

cooling fans. Experience has shown that failures which release the energy stored in 

these rotors may result in engine or structural damage, fires, or injury to occupants. 

The language in this proposal is identical to Secs. 27.1461 and 29.1461.” 

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.” 

Amendment 23-49 and Subsequent 

A proposed revision to NPRM 94-21 clarifies this rule as follows: “This proposal 

would revise paragraph (a) of this section to clarify that the requirements apply to 

high energy rotors included in an auxiliary power unit (APU). Following the 

addition of this section to part 23, the FAA issued a policy message that showed Sec. 

23.1461 was adopted to cover equipment such as APUs and constant speed drives 

that may be installed on small airplanes. The proposed revision of paragraph (a)

will clarify the applicability of this section as identified in that policy material.” 

Amendment 23-49 is additionally clarified by Final Rule, Docket 27806, as follows: 

“This proposal would revise paragraph (a) of this section to clarify that the 

requirements apply to high energy rotors included in an auxiliary power unit (APU). 

One comment was received on this proposal. The JAA and the CAA noted that the 

JAA does not agree that the requirements of this section are applicable to APUs. 

They suggest that the proposed changes to paragraph (a) not be adopted. 

331
AC 23-17B 04/12/05 

In the preamble of the notice, the FAA identified policy issued after this section was 

adopted. That policy indicated that the section was applicable to "equipment such as 

APUs and constant speed drives," but this policy was not widely distributed to all 

FAA offices. The proposal in the notice does not alter the policy applicability, but it 

does clarify the policy. 

Removing the proposed change would not alter the situation. The FAA defines 

"Equipment containing high energy rotors" to include APUs and constant speed 

drives. In cases where rotor containment has been demonstrated by complying with 

JAA-APU or FAA TSO C77a, this compliance will be examined by the FAA office 

responsible for the airplane certification. If it is found that this demonstration also 

meets the requirements of Sec. 23.1461, it will be accepted for the airplanes 

compliance.” 

332