# AC 25-7D CHG 1: Flight Test Guide for Certification of Transport Category Airplanes - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Current - Approved: September 16, 2025 - Last reviewed: September 16, 2025 - Family: AC 25-7 - Cancels: [AC 25-7D](https://openairdocs.org/ac/25-7D) - 14 CFR parts: 25 - 14 CFR sections: [25.903](https://www.ecfr.gov/current/title-14/section-25.903), [25.33](https://www.ecfr.gov/current/title-14/section-25.33), [25.101](https://www.ecfr.gov/current/title-14/section-25.101), [25.105](https://www.ecfr.gov/current/title-14/section-25.105), [25.109](https://www.ecfr.gov/current/title-14/section-25.109), [25.115](https://www.ecfr.gov/current/title-14/section-25.115), [25.21](https://www.ecfr.gov/current/title-14/section-25.21), [25.119](https://www.ecfr.gov/current/title-14/section-25.119), [25.123](https://www.ecfr.gov/current/title-14/section-25.123), [25.125](https://www.ecfr.gov/current/title-14/section-25.125), [25.145](https://www.ecfr.gov/current/title-14/section-25.145), [25.147](https://www.ecfr.gov/current/title-14/section-25.147), [25.173](https://www.ecfr.gov/current/title-14/section-25.173), [25.175](https://www.ecfr.gov/current/title-14/section-25.175), [25.181](https://www.ecfr.gov/current/title-14/section-25.181), [25.231](https://www.ecfr.gov/current/title-14/section-25.231), [25.233](https://www.ecfr.gov/current/title-14/section-25.233), [25.237](https://www.ecfr.gov/current/title-14/section-25.237), [25.239](https://www.ecfr.gov/current/title-14/section-25.239), [25.251](https://www.ecfr.gov/current/title-14/section-25.251), [25.253](https://www.ecfr.gov/current/title-14/section-25.253), [25.255](https://www.ecfr.gov/current/title-14/section-25.255), [25.149](https://www.ecfr.gov/current/title-14/section-25.149), [25.31](https://www.ecfr.gov/current/title-14/section-25.31), [25.107](https://www.ecfr.gov/current/title-14/section-25.107), [25.111](https://www.ecfr.gov/current/title-14/section-25.111), [25.113](https://www.ecfr.gov/current/title-14/section-25.113), [25.117](https://www.ecfr.gov/current/title-14/section-25.117), [25.121](https://www.ecfr.gov/current/title-14/section-25.121), [25.161](https://www.ecfr.gov/current/title-14/section-25.161), [25.177](https://www.ecfr.gov/current/title-14/section-25.177), [25.904](https://www.ecfr.gov/current/title-14/section-25.904), [25.1303](https://www.ecfr.gov/current/title-14/section-25.1303), [25.937](https://www.ecfr.gov/current/title-14/section-25.937), [25.939](https://www.ecfr.gov/current/title-14/section-25.939), [25.959](https://www.ecfr.gov/current/title-14/section-25.959), [25.1351](https://www.ecfr.gov/current/title-14/section-25.1351), [25.1355](https://www.ecfr.gov/current/title-14/section-25.1355), [25.1383](https://www.ecfr.gov/current/title-14/section-25.1383), [25.671](https://www.ecfr.gov/current/title-14/section-25.671), [25.701](https://www.ecfr.gov/current/title-14/section-25.701), [25.729](https://www.ecfr.gov/current/title-14/section-25.729), [25.731](https://www.ecfr.gov/current/title-14/section-25.731), [25.733](https://www.ecfr.gov/current/title-14/section-25.733), [25.735](https://www.ecfr.gov/current/title-14/section-25.735), [25.801](https://www.ecfr.gov/current/title-14/section-25.801), [25.773](https://www.ecfr.gov/current/title-14/section-25.773), [25.1301](https://www.ecfr.gov/current/title-14/section-25.1301), [25.1305](https://www.ecfr.gov/current/title-14/section-25.1305), [25.1309](https://www.ecfr.gov/current/title-14/section-25.1309), [25.1323](https://www.ecfr.gov/current/title-14/section-25.1323), [25.1325](https://www.ecfr.gov/current/title-14/section-25.1325), [25.1329](https://www.ecfr.gov/current/title-14/section-25.1329), [25.1357](https://www.ecfr.gov/current/title-14/section-25.1357), [25.1363](https://www.ecfr.gov/current/title-14/section-25.1363), [25.1381](https://www.ecfr.gov/current/title-14/section-25.1381), [25.1385](https://www.ecfr.gov/current/title-14/section-25.1385), [25.1401](https://www.ecfr.gov/current/title-14/section-25.1401), [25.1431](https://www.ecfr.gov/current/title-14/section-25.1431), [25.1403](https://www.ecfr.gov/current/title-14/section-25.1403), [25.1447](https://www.ecfr.gov/current/title-14/section-25.1447), [25.929](https://www.ecfr.gov/current/title-14/section-25.929), [25.933](https://www.ecfr.gov/current/title-14/section-25.933), [25.943](https://www.ecfr.gov/current/title-14/section-25.943), [25.961](https://www.ecfr.gov/current/title-14/section-25.961), [25.1091](https://www.ecfr.gov/current/title-14/section-25.1091), [25.1093](https://www.ecfr.gov/current/title-14/section-25.1093), [25.1121](https://www.ecfr.gov/current/title-14/section-25.1121), [25.1187](https://www.ecfr.gov/current/title-14/section-25.1187), [25.1197](https://www.ecfr.gov/current/title-14/section-25.1197), [25.1581](https://www.ecfr.gov/current/title-14/section-25.1581) - Pages: 495 - PDF: https://openairdocs.org/files/sha256/31/04/3104934a8e108d885c6020e13f0d42de004808c61f8b34379af65d556341b8c0 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/DRSDOCID117991715620250919115039.0001 - DRS record last changed: November 22, 2025 - Page: https://openairdocs.org/ac/25-7 ## Versions - [AC 25-7D CHG 1](https://openairdocs.org/ac/25-7) (Current), approved Sep 16, 2025, 495 pages - [AC 25-7D](https://openairdocs.org/ac/25-7D) (Historical), approved May 4, 2018, 481 pages - [AC 25-7C](https://openairdocs.org/ac/25-7C) (Historical), approved Oct 16, 2012, 425 pages - [AC 25-7B](https://openairdocs.org/ac/25-7B) (Historical), approved Dec 7, 2011, 410 pages. Change 1 Incorporated - [AC 25-7A](https://openairdocs.org/ac/25-7A) (Historical), approved Mar 31, 1998, 346 pages - [AC 25-7](https://openairdocs.org/ac/25-7?v=59084AF5ED98820B8625733E006D28B2) (Historical), approved Apr 9, 1986, 100 pages ## Text Each page's text from the PDF; open the PDF at a page with #page=N. ### Page 1 Advisory Circular Subject: Flight Test Guide for Certification of Date: 09/16/2025 AC No: 25-7D Transport Category Airplanes Initiated By: AIR-621A Change: 1 1. Purpose. 1.1 This change, advisory circul ar (AC) 25-7D Change 1, incorporates updates and adds or revises guidance for stall in-ground-effect and for tailwind and crosswind capability. 2. Changes. 2.1 Change 1 updates AC 25-7D, Flight Test Guide for Certification of Transport Category Airplanes, dated May 4, 2018. Specifically, Change 1 incorporates the following substantive changes: 1. Flight: General – Section 3.1.10 revises the test tailwind for demonstrating handling qualities from 150% of the proposed limit tailwind to 5 kts above the proposed limit tailwind and the applicable guidance associated with this change. 2. Flight: Performance – Section 4.2 revises the guidance for takeoff performance testing (i.e., 14 CFR 25.107) to raise awareness of ground effects on stall angle-ofattack and maximum lift and to ensure adequate margins to in-ground-effect stall angle-of-attack are maintained during flight testing used to determine takeoff speeds. 3. Flight: Ground and Water Handling Characteristics – Section 9.5 revises the guidance method for determining the magnitude of the maximum demonstrated crosswind to include consideration of gusts. 2.2 The AC change number and the date of the change is shown at the top of each changed page. The change bar in the right or left margin indicates where the change is located. Refer to the Page Control Chart on the next page. ### Page 2 Page Control Chart Remove Pages Dated Insert Pages Dated i through xii 05/04/18 iii through xii 1-1 through 1-4 05/04/18 1-1 through 1-4 3-1, 3-4, 3-9, 3-12 through 3-16 05/04/18 3-1, 3-4, 3-9, 3-12 through 3-16 4-4 through 4-75 05/04/18 4-4 through 4-82 5-1, 5-2, 5-5 through 5-10, 5-15 through 5-20 05/04/18 5-1, 5-2, 5-5 through 5-10,5-15 through 5-20 9-2 through 9-5 05/04/18 9-2 through 9-5 10-3 through 10-15 05/04/18 10-3 through 10-15 28-1 05/04/18 28-1 32-40 05/04/18 32-40 33-20 05/04/18 33-20 C-1 05/04/18 C-1 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 09/16/2025 3. Website Availability. To access this AC electronically, you can go to either of the following sites: FAA Advisory Circulars or the Dynamic Regulatory System. **DANIEL J.** Digitally signed by DANIEL J. ELGAS ELGAS Date: 2025.09.16 16:09:39 -04'00' Daniel J. Elgas Aviation Safety Director, Policy and Standards Division, Aircraft Certification Service i ### Page 3 U.S. Department of Transportation Federal Aviation Administration Subject: Flight Test Guide for Certification of Date: 05/04/2018 AC No: 25-7D Transport Category Airplanes Initiated By: AIR-670 This advisory circular (AC) provides guidance for the flight test evaluation of transport category airplanes. This AC includes flight test methods and procedures to show compliance with the regulations contained in title 14, Code of Federal Regulations (14 CFR) part 25, subpart B, “Flight,” which address airplane performance and handling characteristics. This revision, AC 25-7D, clarifies paragraph 23.2.4, Engine Restart Capability—§ 25.903(e); adds paragraph 34.4, Circuit Protective Devices—§ 25.1357; and revises appendix B, Function and Reliability (F&R) Tests, of this AC. This AC has been re-formatted to use a new paragraph numbering system for improved usability. The first change revises the means of compliance associated with demonstrating the restart capability required by § 25.903(e). The second change adds a paragraph providing guidance for flight test evaluation of compliance with the requirements for circuit protective devices. It is made in response to recommendations from the National Transportation Safety Board. The third change revises appendix B concerning F&R testing. It addresses issues that arose with F&R testing on recent certification programs where the current guidance was unclear. If you have suggestions for improving this AC, you may use the Advisory Circular Feedback form at the end of this AC. /Original signed by Mel Johnson/ Dr. Michael C. Romanowski Director, Policy and Innovation Division Aircraft Certification Service ii ### Page 4 **AC 25-7D CHG 1** **CONTENTS** Chapter 1. Introduction … 1-1 1.1 Purpose … 1-1 1.2 Applicability … 1-1 1.3 Background … 1-1 1.4 Related Material … 1-2 Chapter 2. General … 2-1 2.1 Applicability—§ 25.1. [Reserved] … 2-1 2.2 Special Retroactive Require ments—§ 25.2. [Reserved] … 2-1 Chapter 3. Flight: General … 3-1 3.1 Proof of Compliance—§ 25.21 … 3-1 3.2 Load Distribution Limits—§ 25.23. [Reserved] … 3-15 3.3 Weight Limits and Center of Gravit y Limits—§§ 25.25 and 25.27. [Reserved] … 3-15 3.4 Empty Weight and Corresponding Center of Gravity—§ 25.29. [Reserved] … 3-15 3.5 Removable Ballast—§ 25.31 … 3-15 3.6 Propeller Speed and Pitch Limits—§ 25.33 … 3-16 Chapter 4. Flight: Performance … 4-1 4.1 General—§ 25.101 … 4-1 4.2 Takeoff and Takeoff Speeds—§§ 25.105 and 25.107 … 4-3 4.3 Accelerate-Stop Distance—§ 25.109 … 4-20 4.4 Takeoff Path—§ 25.111 … 4-51 4.5 Takeoff Distance and Takeoff Run—§ 25.113 … 4-59 4.6 Takeoff Flight Path—§ 25.115 … 4-64 4.7 Climb: General—§ 25.117 … 4-66 4.8 Landing Climb: All-Engi nes-Operating—§ 25.119 … 4-66 4.9 Climb: One-Engine-Inoperative—§ 25.121 … 4-68 4.10 En Route Flight Paths—§ 25.123 … 4-71 4.11 Landing—§ 25.125 … 4-72 Chapter 5. Flight: Controllability and Maneuverability … 5-1 5.1 General—§ 25.143 … 5-1 iii ### Page 5 **AC 25-7D CHG 1** CONTENTS (CONTINUED) Paragraph Page 5.2 Longitudinal Control—§ 25.145 … 5-15 5.3 Directional and Late ral Control—§ 25.147 … 5-20 5.4 Minimum Control Speed—§ 25.149 … 5-24 Chapter 6. Flight: Trim … 6-1 6.1 Trim—§ 25.161 … 6-1 6.2 [Reserved.] … 6-1 Chapter 7. Flight: Stability … 7-1 7.1 General—§ 25.171. [Reserved] … 7-1 7.2 Static Longitudinal Stability and Demons tration of Static Longitudinal Stability— §§ 25.173 and 25.175 … 7-1 7.3 Static Directional and Lateral Stability—§ 25.177 … 7-5 7.4 Dynamic Stability—§ 25.181 … 7-10 Chapter 8. Flight: Stalls … 8-1 8.1 Stall Testing … 8-1 Chapter 9. Flight: Ground and Water Handling Characteristics … 9-1 9.1 Part 25 Regulations.................................................................................................... 9- 1 9.2 Longitudinal Stability and Control—§ 25.231 … 9-1 9.3 Directional Stability and Control—§ 25.233 … 9-1 9.4 Taxiing Condition—§ 25.235. [Reserved] … 9-2 9.5 Wind Velocities—§ 25.237 … 9-2 9.6 Spray Characteristics, Control, and Stability on Water—§ 25.239 … 9-6 Chapter 10. Flight: Miscellaneous Requirements … 10-1 10.1 Vibration and Buffeting—§ 25.251 … 10-1 10.2 High Speed Characteristics—§ 25.253 … 10-5 10.3 Out-of-Trim Characteristics—§ 25.255 … 10-11 iv ### Page 6 **AC 25-7D CHG 1** CONTENTS (CONTINUED) Paragraph Page Chapter 11. Structure [Reserved] … 11-1 Chapter 12. Design and Construction: General [Reserved] … 12-1 Chapter 13. Design and Construction: Control Surfaces [Reserved] … 13-1 Chapter 14. Design and Construction: Control Systems … 14-1 14.1 General—§ 25.671 … 14-1 14.2 Flap and Slat Interconnections—§ 25.701 … 14-1 14.3 Takeoff Warning System—§ 25.703. [Reserved] … 14-2 Chapter 15. Design and Construction: Landing Gear … 15-1 15.1 Retracting Mechanism—§ 25.729 … 15-1 15.2 Wheels—§ 25.731 … 15-2 15.3 Tires—§ 25.733 … 15-4 15.4 Brakes—§ 25.735 … 15-4 15.5 Skis—§ 25.737. [Reserved] … 15-18 Chapter 16. Design and Construction: Floats and Hulls [Reserved] … 16-1 Chapter 17. Design and Construction: Personnel and Cargo Accommodations … 17-1 17.1 Pilot Compartment View—§ 25.773 … 17-1 17.2 [Reserved.] … 17-1 Chapter 18. Design and Construction: Emergency Provisions … 18-1 18.1 Ditching—§ 25.801 … 18-1 18.2 Emergency Egress Assist Mean s and Escape Routes—§ 25.810 … 18-1 Chapter 19. Design and Construction: Ventilation and Heating … 19-1 19.1 Ventilation—§ 25.831 … 19-1 19.2 Cabin Ozone Concentration—§ 25.832. [Reserved] … 19-1 19.3 Combustion Heating Systems—§ 25.833. [Reserved] … 19-1 Chapter 20. Design and Construction: Pressurization … 20-1 20.1 Pressurized Cabins—§ 25.841 … 20-1 20.2 Tests for Pressurized Cabins—§ 25.843 … 20-3 v ### Page 7 **AC 25-7D CHG 1** CONTENTS (CONTINUED) Paragraph Page Chapter 21. Design and Construction: Fire Protection [Reserved] … 21-1 Chapter 22. Design and Construction: Miscellaneous [Reserved] … 22-1 Chapter 23. Powerplant: General … 23-1 23.1 Installation—§ 25.901. [Reserved] … 23-1 23.2 Engines—§ 25.903 … 23-1 23.3 Automatic Takeoff Thrust Control System—§ 25.904 … 23-7 23.4 Propellers—§ 25.905. [Reserved] … 23-9 23.5 Propeller Vibration and Fa tigue—§ 25.907. [Reserved] … 23-9 23.6 Propeller Clearance—§ 25.925. [Reserved] … 23-9 23.7 Propeller Deicing—§ 25.929 … 23-9 23.8 Reversing Systems—§ 25.933 … 23-9 23.9 Turbojet Engine Thrust Reverser System Tests—§ 25.934. [Reserved] … 23-12 23.10 Turbopropeller-Drag Limiting Systems—§ 25.937 … 23-12 23.11 Turbine Engine Operating Characteristics—§ 25.939 … 23-14 23.12 Inlet, Engine, and Exhaust Comp atibility—§ 25.941. [Reserved] … 23-15 23.13 Negative Acceleration—§ 25.943 … 23-15 23.14 Thrust or Power Augmentation System—§ 25.945. [Reserved] … 23-16 Chapter 24. Powerplant: Fuel System … 24-1 24.1 Unusable Fuel Supply—§ 25.959 … 24-1 24.2 Fuel System Hot Weather Operation—§ 25.961 … 24-3 24.3 Fuel Tank Vents and Carburetor Vapor Vents—§ 25.975 … 24-5 Chapter 25. Powerplant: Fuel System Components … 25-1 25.1 Fuel Jettisoning System—§ 25.1001 … 25-1 25.2 [Reserved.] … 25-3 Chapter 26. Powerplant: Oil System [Reserved] … 26-1 Chapter 27. Powerplant: Cooling … 27-1 27.1 Cooling Test Procedures—§ 25.1045 … 27-1 27.2 [Reserved.] … 27-8 vi ### Page 8 **AC 25-7D CHG 1** CONTENTS (CONTINUED) Paragraph Page Chapter 28. Powerplant: Induction System … 28-1 28.1 Air Induction—§ 25.1091 … 28-1 28.2 Induction System Ici ng Protection—§ 25.1093 … 28-3 Chapter 29. Powerplant: Exhaust System … 29-1 29.1 General—§ 25.1121 … 29-1 29.2 [Reserved.] … 29-1 Chapter 30. Powerplant: Controls and Accessories [Reserved] … 30-1 Chapter 31. Powerplant: Fire Protection … 31-1 31.1 Drainage and Ventilation of Fire Zones—§ 25.1187 … 31-1 31.2 Fire Extinguishing Systems—§ 25.1197 … 31-1 Chapter 32. Equipment: General … 32-1 32.1 Equipment: Function an d Installation—§ 25.1301 … 32-1 32.2 Flight and Navigation Instruments—§ 25.1303 … 32-39 32.3 Powerplant Instruments—§ 25.1305. [Reserved] … 32-40 32.4 Miscellaneous Equipment—§ 25.1307. [Reserved] … 32-40 32.5 Equipment, Systems, an d Installations—§ 25.1309 … 32-40 Chapter 33. Equipment: Instruments Installation … 33-1 33.1 Arrangement and Visibility—§ 25.1321. [Reserved] … 33-1 33.2 Warning, Caution, and Advisory Lights—§ 25.1322. [Reserved] … 33-1 33.3 Airspeed Indicating System—§ 25.1323 … 33-1 33.4 Static Pressure Systems—§ 25.1325(d) and (e) … 33-4 33.5 Pitot Heat Indication Syst ems—§ 25.1326. [Reserved] … 33-5 33.6 Magnetic Direction Indicato r—§ 25.1327. [Reserved] … 33-5 33.7 Flight Guidance System—§ 25.1329 … 33-5 Chapter 34. Equipment: Electrical Systems and Equipment … 34-1 34.1 General—§ 25.1351 … 34-1 34.2 Electrical Equipment and Inst allations—§ 25.1353. [Reserved] … 34-4 34.3 Distribution System—§ 25.1355 … 34-4 vii ### Page 9 **AC 25-7D CHG 1** CONTENTS (CONTINUED) Paragraph Page 34.4 Circuit Protective Devices—§ 25.1357 … 34-4 34.5 Electrical System Tests—§ 25.1363 … 34-5 Chapter 35. Equipment: Lights … 35-1 35.1 Instrument Lights—§ 25.1381 … 35-1 35.2 Landing Lights—§ 25.1383 … 35-1 35.3 Position Light System Installation—§ 25.1385 … 35-1 35.4 Anti-Collision Light System—§ 25.1401 … 35-2 35.5 Wing Icing Detection Lights—§ 25.1403 … 35-3 Chapter 36. Safety Equipment [Reserved] … 36-1 Chapter 37. Miscellaneous Equipment … 37-1 37.1 Electronic Equipment—§ 25.1431 … 37-1 37.2 Equipment Standards for Oxyge n Dispensing Units—§ 25.1447 … 37-2 Chapter 38. Operating Limitations and Information: General [Reserved] … 38-1 Chapter 39. Operating Limitations and Information: Operating Limitations [Reserved] … 39-1 Chapter 40. Operating Limitations and Information: Markings and Placards [Reserved] … 40-1 Chapter 41. Operating Limitations and Information: Airplane Flight Manual … 41-1 41.1 General—§ 25.1581 … 41-1 41.2 [Reserved.] … 41-1 Chapter 42. Airworthiness: Miscellaneous Items … 42-1 42.1 Design and Function of Artificial Stall Wa rning and Identification Systems … 42-1 42.2 Reduced and Derated Power or Th rust Takeoff Operations … 42-5 42.3 Runway Gradients Greater than ±2 Percent … 42-6 42.4 Criteria for Approval of Steep Approach to Landing … 42-8 42.5 Takeoff and Landing on Unpaved Runways … 42-12 42.6 Accounting for Performance Effects of Minor Design Changes and Configuration Deviation List Items … 42-15 42.7 Configuration Deviation List … 42-16 42.8 Spare Engine Pod … 42-19 viii ### Page 10 **AC 25-7D CHG 1** CONTENTS (CONTINUED) Paragraph Page 42.9 Authorization for Ferry Flight wi th One Engine Inoperative—§ 91.611 … 42-19 42.10 Instrument Landing System Weather Minima … 42-20 42.11 Takeoff Performance Credit for Alternate Forward Center of Gravity Limits … 42-21 42.12 Airplane Backing Usi ng Reverse Thrust … 42-23 appendix A. Acronyms, Abbreviations, Symbols, and Definitions … A-1 appendix B. Function and Reliability (F&R) Tests … B-1 appendix C. Historical Development of Accelerate-Stop Time Delays … C-1 appendix D. History of Jet Transport Performance Standards … D-1 appendix E. FAA Handling Qualities Rating Method … E-1 appendix F. Correction of Air Minimum Control Speed to Standard Conditions … F-1 appendix G. Rudder Pedal Force-Limited Air Minimum Control Speed … G-1 ix ### Page 11 **AC 25-7D CHG 1** **CONTENTS (CONTINUED)** FIGURES Number Page Figure 3-1. Equivalent Weight Extrapolation … 3-6 Figure 3-2. Wind Profile Variation … 3-9 Figure 4-1. Accelerate-Stop Time Delays … 4-29 Figure 4-2. Accelerate-Stop Speed versus Distance … 4-31 Figure 4-3. Anti-Skid System Response Characteristics: On-Off System on Dry Runway … 4-36 Figure 4-4. Anti-Skid System Response Characteristics: On-Off System on Wet Runway … 4-36 Figure 4-5. Anti-Skid System Response Characteristics: Quasi-Modulating System on Dry Runway … 4-37 Figure 4-6 Anti-Skid System Response Characteristics: Quasi-Modulating System on Wet Runway … 4-38 Figure 4-7. Anti-Skid System Response Characteristics: Fully Modulating System … 4-39 Figure 4-8. Instantaneous Brake Force and Peak Brake Force … 4-43 Figure 4-9. Anti-Skid Efficiency–Wheel Slip Relationship … 4-45 Figure 4-10. Substantiation of the Optimal Slip Value … 4-46 Figure 4-11. Takeoff Distance on a Dry Runway: Critical Engine Fails at VEF … 4-60 Figure 4-12. Takeoff Distance: All-Engines-Operating … 4-60 Figure 4-13. Takeoff Distance on a Wet Runway: Critical Engine Fails at VEF … 4-61 Figure 4-14. Takeoff Run: Critical Engine Fails at VEF … 4-62 Figure 4-15. Takeoff Run: All-Engines-Operating … 4-63 Figure 4-16. Clearway Profiles … 4-63 Figure 4-17. Takeoff Segments and Nomenclature … 4-65 Figure 4-18. Net Takeoff Flight Path … 4-66 Figure 4-19. Landing Time Delays … 4-79 Figure 5-1. Sample Pitch Tracking Task … 5-10 Figure 7-1. Longitudinal Static Stability … 7-4 Figure 7-2. Local Reversal … 7-5 Figure 8-1. Thrust Effect on Stall Speed … 8-5 Figure 8-2. CLMAX and Load Factor … 8-7 Figure 8-3. CLMAX versus Weight and Flap Setting … 8-8 x ### Page 12 **AC 25-7D CHG 1** **CONTENTS (CONTINUED)** FIGURES Number Page Figure 8-4. Stall Speed versus Weight and Flap Setting … 8-10 Figure 10-1. Maneuvering Characteristics at Speeds up to VMO/MMO … 10-4 Figure 10-2. Mistrimmed Maneuvering Characteristics: Speeds Between VFC/MFC and VDF/MDF . … 10-13 Figure 33-1. Altitude Loss versus Altitude at Go-Around Mode Initiation … 33-16 Figure 33-2. Deviation Profile Method … 33-26 Figure 33-3. Height Loss Method … 33-27 Figure C-1. Accelerate-Stop Time Delays (Pre-Amendment 25-42) … C-2 Figure C-2. Accelerate-Stop Time Delays (Amendment 25-42 through Amendment 25-91) ... C-3 Figure E-1. Overall HQRM Process … E-2 Figure E-2. Sample Tasks for Evaluating Airplane Handling Qualities … E-4 Figure E-3. Probability of Occurrence Guidelines … E-5 Figure E-4. Flaps UP Flight Envelopes … E-7 Figure E-5. Flaps DOWN Flight Envelopes … E-8 Figure E-6. Combining Values … E-9 Figure E-7. Probability Guidelines to Determine HQ Requirements … E-10 Figure F-1. Two-Engine Business Jet … F-4 Figure F-2. Four-Engine Transport … F-4 Figure F-3. Yawing Moment—Engine and Airframe … F-5 Figure F-4. Yawing Moment—Engine and Airframe at Altitude of 3,000 Feet (Test Day) … F-7 Figure G-1. Rudder Force Versus WsinΦ … G-2 Figure G-2. Plot to Determine VMCA … G-4 Figure G-3. Plot to Determine VMCA for an Amendment 25-42 Airplane … G-5 Figure G-4. Plot to Determine VMCA for Derate Thrust … G-6 xi ### Page 13 **AC 25-7D CHG 1** CONTENTS (CONTINUED) TABLES Number Page Table 3-1. Weight Tolerance Limits … 3-4 Table 3-2. Test Parameters that Normally can be Corrected … 3-8 Table 5-1. PIO Rating Criteria and Comparison to MIL Standard … 5-11 Table 5-2. Example of Acceptable HQ Rating for PIO Tendencies … 5-12 Table 32-1. Radio Line-of-Sight Distance versus Flight Level or Altitude … 32-5 Table 32-2. Marker Beacon System … 32-12 Table 32-3. Classification of LF/MF Radio Beacons in U.S. National Service … 32-14 Table 32-4. Types of Aural Warnings … 32-33 Table A-1. Terms and Definitions … A-1 Table A-2. Symbols and Definitions … A-10 Table E-1. Comparison of Handling Qualities Ratings … E-2 Table E-2. Minimum HQ Requirements … E-11 Table F-1. Example Data for Typical Business Jet Engine … F-5 Table F-2. Example Data for Typical Business Jet Engine at Altitude of 3,000 Feet … F-7 xii ### Page 14 **AC 25-7D CHG 1 CHAPTER 1. INTRODUCTION** 1.1 Purpose. 1.1.1 This AC provides updated guidance for the flight test evaluation of transport category airplanes. These guidelines provide an acceptable means of demonstrating compliance with the pertinent regulations of Title 14, Code of Federal Regulations (14 CFR) part 25. The methods and procedures described herein have evolved through many years of flight testing of transport category airplanes and, as such, represent current certification practice. 1.1.2 See appendix A for a list of acronyms and abbreviations used in this AC. 1.2 Applicability. 1.2.1 The guidance provided in this document is directed to airplane manufacturers, modifiers, foreign regulatory authorities, and Federal Aviation Administration (FAA) certification engineers, flight test pilots, and designees. 1.2.2 This is a guidance document. Its content is not legally binding in its own right and will not be relied upon by the Department as a separate basis for affirmative enforcement action or other administrative penalty. Conformity with the guidance document is voluntary only. Nonconformity will not affect rights and obligations under existing statutes and regulations. 1.3 Background. 1.3.1 Since AC 25-7 was released on April 9, 1986, it has been the primary source of guidance for flight test methods and procedures to show compliance with the regulations contained in subpart B of part 25, which address airplane performance and handling characteristics. AC 25-7 has been revised several times to reflect changes in the part 25 regulatory requirements, changes in guidance and policy, and advances in technology. 1.3.2 The first revision, AC 25-7A, updated the original AC to incorporate the policy and guidance material applicable to all sections of part 25, not just subpart B. The material related to regulations outside of subpart B superseded that contained in Order 8110.8, Engineering Flight Test Guide for Transport Category Airplanes, which was cancelled when AC 25-7A was issued. 1.3.3 Change 1 to AC 25-7A added acceptable means of compliance for the regulatory changes associated with amendments 25-92 and 25-98 to part 25. 1.3.4 AC 25-7B added acceptable means of compliance for the regulatory changes associated with amendments 25-108, 25-109, and 25-115 to part 25, and revised guidance for 1-1 ### Page 15 **AC 25-7D CHG 1** expanding takeoff and landing data for airport elevations higher than those at which flight testing was conducted. Means of compliance associated with flight in icing conditions were removed as this material is now contained in AC 25-25A, Performance and Handling Characteristics in Icing Conditions, dated October 27, 2014. 1.3.5 Change 1 to AC 25-7B added acceptable means of compliance for the regulatory changes associated with amendment 25-135. 1.3.6 AC 25-7C reduced the number of differences between the FAA and European Aviation Safety Agency flight test guides, provided acceptable means of compliance for the regulatory changes associated with amendments 25-107, 25-109, 25-113, 25-115, 25-119 and 25-123 to part 25, and included changes responding to safety recommendations from the FAA and National Transportation Safety Board. 1.3.7 AC 25-7D clarified paragraph 23.2.4, Engine Restart Capability—§ 25.903(e); added paragraph 34.4, Circuit Protective Devices—§ 25.1357; and revised appendix B, Function and Reliability (F&R) Tests, of this AC. The AC was reformatted to use a new paragraph numbering system for improved usability. 1.3.8 Change 1 to AC 25-7D incorporates updates and adds or revises guidance for stall inground-effect and for tailwind and crosswind capability in response to several National Transportation Safety Board safety recommendations. 1.4 Related Material. 1.4.1 Orders. The following FAA orders are related to the guidance in this AC. The latest version of each order at the time of publication of this AC is identified below. If any order is revised after publication of this AC, you should refer to the latest version for guidance, which can be downloaded from the internet at https://www.faa.gov/regulations_policies/orders_notices/.  Order 8100.5E, Aircraft Certification Service – Organizational Structure and Functions, dated April 9, 2023.  Order 8110.4C, with Change 7, Type Certification, dated October 20, 2023. 1.4.2 Advisory Circulars. The following ACs are related to the guidance in this AC. The latest version of each AC referenced in this document is available on the FAA website at FAA Advisory Circulars and on the Dynamic Regulatory System.  AC 20-73A, Aircraft Ice Protection, dated August 16, 2006.  AC 20-124, Water Ingestion Testing for Turbine Powered Airplanes, dated September 30, 1985. 1-2 ### Page 16 AC 25-7D CHG 1  AC 20-131A, Airworthiness Approval of Traffic Alert and Collision Avoidance Systems (TCAS II) and Mode S Transponders, dated March 29, 1993.  AC 20-138D, with Change 2, Airworthiness Approval of Positioning and Navigation Systems, dated April 7, 2016.  AC 20-147A, Turbojet, Turboprop, Turboshaft, and Turbofan Engine Induction System Icing and Ice Ingestion, dated October 22, 2014.  AC 20-161, Aircraft Onboard Weight and Balance Systems, dated April 11, 2008.  AC 20-168, Certification Guidance for Installation of Non-Essential, Non-Required Aircraft Cabin Systems & Equipment (CS&E), dated July 22, 2010,  AC 21-29D, Detecting and Reporting Suspected Unapproved Parts, dated July 12, 2016.  AC 21.101-1B, Establishing the Certification Basis of Changed Aeronautical Products, March 11, 2016.  AC 25-9A, Smoke Detection, Penetration, and Evacuation Tests and Related Flight Manual Emergency Procedures, dated January 6, 1994.  AC 25-11B, Electronic Flight Displays, dated October 7, 2014.  AC 25-12, Airworthiness Criteria for the Approval of Airborne Windshear Warning Systems in Transport Category Airplanes, dated November 2, 1987.  AC 25-13, Reduced and Derated Takeoff Thrust (Power) Procedures, dated May 4, 1988.  AC 25-15, Approval of Flight Management Systems in Transport Category Airplanes, dated November 20, 1989.  AC 25-17A, with Change 1, Transport Airplane Cabin Interiors Crashworthiness Handbook, dated May 24, 2016.  AC 25-20, Pressurization, Ventilation, and Oxygen Systems Assessment for Subsonic Flight Including High Altitude Operations dated September 10, 1996.  AC 25-22, Certification of Transport Airplane Mechanical Systems, dated March 14, 2000.  AC 25-23, Airworthiness Criteria for Installation Approval of a Terrain Awareness and Warning System (TAWS) for Part 25 Airplanes, dated May 22, 2000.  AC 25-25A, Performance and Handling Characteristics in Icing Conditions, dated October 27, 2014.  AC 25.735-1, Brakes and Braking Systems Certification Tests and Analysis, dated April 10, 2002.  AC 25.773-1, Pilot Compartment View Design Considerations, dated January 8, 1993. 1-3 ### Page 17 AC 25-7D CHG 1  AC 25.939-1, Evaluating Turbine Engine Operating Characteristics, dated March 19, 1986.  AC 25.1309-1B, System Design and Analysis, dated August 30, 2024.  AC 25.1329-1C, with Change 1, Approval of Flight Guidance Systems, dated May 24, 2016.  AC 25.1357-1A, Circuit Protective Devices, dated October 22, 2007.  AC 25.1581-1, with Change 1, Airplane Flight Manual, dated October 16, 2012.  AC 43.13-1B, with Change 1, Acceptable Methods, Techniques, and Practices— Aircraft Inspection and Repair, dated September 27, 2001.  AC 90-100A, with Change 2, U.S. Terminal and En Route Area Navigation (RNAV) Operations, dated April 14, 2015.  AC 90-101A, with Change 1, Approval Guidance for RNP Procedures with AR, dated February 9, 2016.  AC 90-105A, Approval Guidance for RNP Operations and Barometric Vertical Navigation in the U.S. National Airspace System and in Oceanic and Remote Continental Airspace, dated March 7, 2016.  AC 91-79A, with Change 1, Mitigating the Risks of a Runway Overrun Upon Landing, dated April 28, 2016.  AC 120-28D, Criteria for Approval of Category III Weather Minima for Takeoff, Landing, and Rollout, dated July 13, 1999.  AC 120-29A, Criteria for Approval of Category I and Category II Weather Minima for Approach, dated August 12, 2002.  AC 150/5320-12C, with Change 8, Measurement, Construction, and Maintenance of Skid-Resistant Airport Pavement Surfaces, dated February 7, 2007. 1-4 ### Page 18 **CHAPTER** **2. GENERAL** 2.1 Applicability—§ 25.1. [Reserved] 2.2 Special Retroactive Requirements—§ 25.2. [Reserved] 2-1 ### Page 19 AC 25-7D CHG 1 CHAPTER 3. FLIGHT: GENERAL 3.1 Proof of Compliance—§ 25.21. 3.1.1 Explanation. To provide the necessary guidelines for the flight test evaluation of transport category airplanes, without producing a cumbersome document, this AC assumes a conventional transport airplane configuration. In general, a conventional airplane configuration is one with distinct wing and fuselage elements that are joined together, aft-mounted horizontal and vertical stabilizers that are attached to the fuselage, and propulsion provided either by turbojet/turbofan engines that do not provide any significant increase in lift due to their operation or engine-driven propellers. The effects of non-conventional airplane configurations (e.g., blown flaps) on the compliance methods should be evaluated and determined based on the intent of the guidelines presented for conventional airplane configurations. 3.1.2 Section 25.21(a)—Proof of Compliance. 3.1.2.1 The burden of showing compliance with the flight requirements for an airworthiness certificate or a type certificate rests with the applicant. The applicant should, at their own expense and risk, conduct such official flight tests as required by the FAA to demonstrate compliance with the applicable requirements. During the certification process, the applicant should make available the airplane, as well as all of the personnel and equipment necessary to obtain and process the required data. 3.1.2.2 If the airplane flight characteristics or the required flight data are affected by weight and/or center of gravity (CG), the compliance data must be presented for the most critical weight and CG position per § 25.21(a). Unless the applicant shows that the allowable CG travel in one or more axes (e.g., lateral fuel imbalance) has a negligible effect on compliance with the airworthiness requirements, the applicant must substantiate compliance at the critical CG. 3.1.2.3 The gross weight and CG tolerances specified in paragraphs 3.1.4.3 and 3.1.4.5 are test tolerances and are not intended to allow compliance to be shown at less than critical conditions. 3.1.2.4 Section 21.35(a)(3) requires that the test airplane be in conformity with its type design specifications. This means that the test airplane must be in conformity with its type design specification as it relates to the particular test being conducted. Any deviation from conformity must be clearly shown to be of no consequence to the particular test being conducted. For example, if the slip resistant escape surface required by § 25.810(c) is not installed when conducting airplane performance and flight characteristics 3-1 ### Page 20 tests, the applicant must show that its presence would have no effect on measured airplane performance and flight characteristics. 3.1.2.5 Section 21.35(b)(2) requires the applicant to conduct sufficient flight testing the FAA finds necessary to determine whether there is reasonable assurance that the airplane, its components, and its equipment are reliable and function properly. Appendix B to this AC provides guidance for showing compliance with this requirement. 3.1.2.6 Acceptable Use of Simulation in Lieu of Flight Testing. It is difficult to establish guidance for using simulation in lieu of flight testing that applies in all situations. However, the following general principles can be used as guidance for determining the acceptability for using simulation in lieu of flight testing: 3.1.2.6.1 In general, flight test demonstrations are the preferred method to show compliance. 3.1.2.6.2 Simulation may be an acceptable alternative to flight demonstrations in certain situations, such as the following: 1. A flight demonstration would be too risky even after attempts are taken to mitigate these risks (e.g., by mock takeoffs/landings in the air at a safe altitude). 2. The required environmental or airplane conditions are too difficult to attain, such as (1) validation of system safety analyses failure cases involving high crosswinds; (2) development of crosswind guidance for slippery runway operations; and (3) conditions involving minimum allowable weight where the minimum allowable weight cannot be achieved because of the weight of required test equipment. In case (3), simulation data can be used to supplement flight test data obtained at the minimum practicable test weight. 3. The simulation is used to augment a reasonably broad flight test program; or 4. The simulation is used to demonstrate repeatability, or to demonstrate performance of a specific scenario for a range of pilots. 3.1.2.6.3 Simulation Criteria. If it is agreed that a simulation will be used to establish compliance, then the simulation should meet the following criteria in order to be acceptable for showing compliance with the performance and handling qualities requirements: 1. The simulation should be of a type and fidelity that is appropriate for the task. For example, is motion or an exterior view needed, or is the fidelity or customizability of an engineering simulator needed? 3-2 ### Page 21 2. The simulation should be suitably validated by flight test data for the conditions of interest. This does not mean that there must be flight test data at the exact conditions of interest. The reason simulation is being used may be that it is too difficult or risky to obtain flight test data at the conditions of interest. The level of substantiation of the simulator to flight correlation should be commensurate with the level of compliance (i.e., the closer the case is to being non-compliant, the higher the required fidelity of the simulation). 3. The simulation should be conducted in a manner appropriate to the case and conditions of interest. If closed-loop responses are important, the simulation should be piloted by a human pilot. For piloted simulations, the controls/displays and cues should be substantially equivalent to what would be available in the real airplane (unless it is determined that not doing so would provide added conservatism). 3.1.3 Section 25.21(c)—Proof of Compliance (Altitude Effect on Flight Characteristics). 3.1.3.1 Any of the flying qualities affected by altitude, including controllability, stability, trim, and stall characteristics, must be investigated at the most adverse altitude conditions approved for operations. 3.1.3.2 Consideration should be given in the test program to any aerodynamic control system changes that occur with changes in altitude (e.g., maximum control surface displacement or auto slats that may be inhibited by Mach number above a specific altitude). 3.1.4 Section 25.21(d)—Proof of Compliance: Flight Test Tolerances. 3.1.4.1 To allow for variations from precise test values, acceptable tolerances during flight testing must be maintained. The purpose of these tolerances is to allow for small variations in flight test values of certain variables from the targeted value. They are not intended for compliance tests to be planned for other than the critical condition, nor are they to be considered as an allowable measurement error. 3.1.4.2 Where variation in the parameter for which a tolerance is allowed will have an effect on the results of the test, the results should be corrected to the most critical value of that parameter within the approved operating envelope. If such a correction is impossible or impractical, the average test conditions should assure that the measured characteristics represent the actual critical value. 3.1.4.3 Weight Limits. 3.1.4.3.1 Table 3 1 below presents weight tolerances that have been found acceptable for the specified flight tests. Many flight tests need to be 3-3 ### Page 22 AC 25-7D CHG 1 conducted at or very near the maximum operating weight for the airplane configuration, particularly those tests used to establish airplane flight manual (AFM) performance information. As noted in paragraph 3.1.4.1 above, the purpose of the test tolerances is to allow for variations in flight test values, not to routinely schedule tests at less than critical weight conditions or to allow compliance to be shown at less than the critical weight condition. In addition, the tolerances can be used to help determine when to interrupt a series of test conditions in order to refuel the airplane if necessary to remain within the acceptable weight tolerance. Table 3-1. Weight Tolerance Limits Weight Tolerance Limit Flight Test Condition ±5% ±10% Stall Speeds X Stall Characteristics X All Other Flight Characteristics X Climb Performance X Takeoff Flight Paths X Landing Braking Distance X Landing Air Distance X Takeoff Distance and Speed X Accelerate-Stop Distance X Maximum Energy RTOs X Minimum Unstick Speed X Minimum Control Speed X Note: A -5 percent tolerance limit means that the weight for the particular test may be up to 5 percent less than the test target value. A +5 percent tolerance limit means that the weight for the particular test may be up to 5 percent higher than the test target value. 3.1.4.3.2 It can be difficult or impossible to conduct testing at the airplane’s minimum allowable weight with an airplane configured for conducting a flight test program. If the minimum weight cannot be obtained (within the specified tolerance limit) and compliance at the minimum weight cannot be clearly deduced from the results at the tested weight, the testing should be conducted on a production airplane (or other airplane on which the minimum weight can be obtained). If the instrumentation or equipment. 3-4 ### Page 23 needed to conduct safe testing cannot be installed on the production airplane configuration, or the weight of such instrumentation still prevents the minimum weight from being obtained, consider the use of simulation to extend the results obtained at the minimum practical test weight. (See paragraph 3.1.2.6 of this AC.) 3.1.4.3.3 For follow-on airplane certification programs involving an increase in the maximum allowable gross weight, the test weight limits of table 3-1 have been applied as extrapolation limits on the original test data in order to minimize additional testing. For the test weight tolerance limits to be applied in this manner, the original test data must be from an existing certificated database for an aerodynamically similar model of the same airplane type. The tolerance limit should be applied to the maximum weight at which the original testing was conducted, not to the maximum certified weight. 3.1.4.3.4 Equivalent Weight Extrapolation Limits. For follow-on airplane certification programs where it is desired to increase a maximum operating weight based on existing certified performance parameters that have weight as one of their independent terms, those parameters should be examined for equivalent compliance with the weight tolerance limits of table 3-1. An example would be the reduction of an airplane’s landing flap position, to one approved on a similar model of the same airplane type, which would incur an increase in landing speeds and brake energy, relative to the original certificated landing flap, at any given weight. The brake energy, at the maximum certificated landing weight, should be calculated for the reduced landing flap. This brake energy should account for the increased landing speeds and reduced aerodynamic drag associated with the reduced flap setting. It should then be determined what equivalent gross weight would have rendered that brake energy with the original landing flap. (See figure 3-1 below for an example of how this can be done.) If the resulting equivalent gross weight does not exceed the certificated maximum landing weight by more than the five percent weight extrapolation limit specified in table 3-1, the reduced flap certification may be eligible for a reduced flight test program (e.g., limited to stall speed verification, handling characteristics, and a qualitative landing demonstration). Further limitations may be imposed by the criteria of technical standard order (TSO) C135a, Transport Airplane Wheels and Wheel and Brake Assemblies, dated July 1, 2009. 3-5 ### Page 24 Figure 3-1. Equivalent Weight Extrapolation 3.1.4.4 Wind Limits. A wind velocity limit of 10 knots (from any direction) or 0.11 VSR1 (whichever is lower) is considered the maximum acceptable for obtaining valid takeoff and landing flight test data. Takeoff and landing performance data obtained under runway wind conditions greater than 5 knots may be inconsistent and unreliable because winds of that magnitude are likely to be unsteady. However, performance data obtained with winds between 5 and 10 knots should not necessarily be discarded. Their validity should be checked against data obtained in conditions with lesser winds. Wind velocity should be measured at the height of the wing mean aerodynamic chord (MAC), as determined with the airplane in a static ground attitude. When measuring test wind velocity at the wing MAC height, a height of six feet above the ground should be considered as a minimum measurement height to avoid possible measurement inaccuracies due to surface interference. 3.1.4.5 CG Limits. A test tolerance of ±7 percent of the total CG range is intended to allow for inflight CG movement. This tolerance is only acceptable when the test data scatter is on both sides of the limiting CG or when adjusting the data from the test CG to the limit CG is acceptable. If compliance with a requirement is marginal at a test condition that is inside of the CG limits, the test should be repeated at the CG limits. 3.1.4.6 Airspeed Limits. Normally, tests conducted within 3 percent or 3 knots (whichever is the higher) of the desired test speed are considered acceptable. 3-6 ### Page 25 3.1.4.7 Thrust/Power Limits. Thrust critical tests, such as minimum control speeds, should be conducted at the highest thrust (or power) allowable on the engine given the constraints of temperature and altitude. It is then permitted to calculate further corrections to allow extrapolation of data to cover the entire operating envelope. These thrust (or power) corrections should be limited to 5 percent of test day thrust (or power) unless a detailed analysis is performed. 3.1.4.8 It is not the purpose of these tolerances to allow flights at values in excess of those authorized in the type design. If such flights are to be conducted, adequate structural substantiation for the flight conditions should be available. These flights should always be conducted under controlled conditions and with the flight test crew’s full knowledge of the situation. Examples of such flights are: 3.1.4.8.1 Takeoff at greater than maximum takeoff weight to reach a test area at the maximum takeoff weight. 3.1.4.8.2 Landing at greater than maximum landing weights during the course of conducting takeoff tests. 3.1.4.8.3 Flights to obtain data for future approvals beyond that substantiated for the initial type design. 3.1.4.8.4 Table 3-2 below indicates the cases for which corrections are normally allowed. Any corrections to flight test data should be made by methods that are agreed to by the FAA. 3-7 ### Page 26 Table 3-2. Test Parameters that Normally can be Corrected Flight Test Condition Correctable Parameters Weight CG Airspeed Altitude Power/ Thrust Wind Airspeed calibration X --- --- --- --- --- Stall speeds X X --- --- X --- Climb performance X X X X X --- Landing performance X --- X X --- X Takeoff performance X X --- X X X Accelerate-stop performance X X --- X X X Minimum control speed --- --- --- --- X --- Minimum unstick speed X X X --- X --- Buffet boundary X X --- X --- --- 3.1.4.9 All instrumentation used in the flight test program should be appropriately calibrated and acceptable to the FAA test team. 3.1.5 Section 25.21(f)—Proof of Compliance: Wind Measurement and Corrections. 3.1.5.1 The relationship between the wind measured at one height and the corresponding wind at another height may be obtained by the following equation: Where: 3.1.5.2 This equation is presented graphically below. Values of H less than 5 feet should not be used in this relationship. 3-8 ### Page 27 AC 25-7D CHG 1 Figure 3-2. Wind Profile Variation 0 10 20 30 40 50 60 70 HEIGHT ABOVE FIELD ELEVATION - FEET 10 meters 0 1 0 2 0 3 0 4 0 5 0 6 0 **WIND VELOCITY - KNOTS** 3.1.6 Wind Profile Variation for Test Data. The performance data of airplanes should be obtained in such a manner that the effect of wind on the test data may be determined. The test wind velocity should be corrected from the recorded height above the test surface to the height of the airplane wing mean aerodynamic chord. If the wind profile variation is not measured, the variation may be calculated using the equation in paragraph 3.1.5 above. The following examples are methods of handling wind profile variation data. Other methods have also been found acceptable (see paragraph 9.5.2.3). Example: Test Data Given:  Height of mean aerodynamic chord with airplane on 8.0 feet surface  Height of wind measurement 6.0 feet  Measured wind velocity 4.8 knots Results:  Test wind velocity with airplane 50 feet above landing 4.8((50 + 8)/6)1/7 = 6.6 surface knots  Test wind velocity with airplane 35 feet above takeoff 4.8((35 + 8)/6)1/7 = 6.4 surface knots  Test wind velocity with airplane on surface 4.8(8/6)1/7 = 5.0 knots 3-9 ### Page 28 3.1.7 Wind Profile Variation for AFM Data. When expanding the data to the AFM conditions, the result should include the effective velocity, at the airplane’s wing mean aerodynamic chord, which corresponds to the wind condition as measured at 10 meters (32.81 feet) above the takeoff surface and corrected for the wind factors of § 25.105(d)(1). Example: AFM Data Given:  Height of mean aerodynamic chord with 8.0 feet airplane on surface  Reported headwind at 10 meters 40.0 knots  Section 25.105(d)(1) wind factor 0.5 Results:  Factored wind velocity with airplane 50 feet (0.5)(40)((50 + 8)/32.81)1/7 = 21.7 knots above landing surface  Factored wind velocity with airplane 35 feet (0.5)(40)((35 + 8)/32.81)1/7 = 20.8 knots above takeoff surface  Factored wind velocity with airplane on (0.5)(40)(8/32.81)1/7 = 16.3 knots surface 3.1.8 Airplane Airspeed Variation Due to Wind Profile Variation Combined with Speed Changes Due to Airplane Dynamic Performance. In the reduction of test data and in the expansion of such data to AFM conditions, the increase or decrease of speed due to the dynamic effect of the forces on the airplane are shown only by the change in ground speed. These changes in ground speed may be generalized either as speed increments or speed ratios. The changes in airspeed due to wind profile variation are superimposed on these speed changes. Example: Determination of True Airspeed from Ground Speed—Takeoff Test Data Given:  Ground speed at liftoff, VLOF 139.0 knots  Ground speed at 35 feet above takeoff surface 140.6 knots  Speed change due to airplane dynamic performance 1.6 knots  Test headwind at liftoff 5.0 knots  Test headwind with airplane 35 feet above takeoff surface 6.4 knots Results:  True airspeed at liftoff, VLOF 139.0 + 5.0 = 144.0 knots  True airspeed at 35 feet above takeoff surface 140.6 + 6.4 = 147.0 knots 3-10 ### Page 29 Example: Determination of Rotation Speed from True Airspeed at 35-Foot Height—AFM Data Given:  Factored headwind at liftoff 16.3 knots  Factored headwind with airplane 35 feet above takeoff surface 20.8 knots  Ground speed change, (V35 – VLOF) 1.6 knots  Ground speed change, (VLOF – VR) 0.5 knots  True airspeed required at 35 feet 150.0 knots Results:  Ground speed required at 35 feet 150 - 20.8 = 129.2 knots  Ground speed at liftoff 129.2 - 1.6 = 127.6 knots  True airspeed at liftoff 127.6 + 16.3 = 143.9 knots  Ground speed at rotation 127.6 - 0.5 = 127.1 knots  True airspeed at rotation (for AFM speed and distances) 127.1+16.3 = 143.4 knots Example: Landing—AFM Data Given:  Factored headwind with airplane 50 feet above landing surface 21.7 knots  Factored headwind with airplane on landing surface 16.3 knots  Ground speed change for 50 feet to touchdown (V50 – VTD) 4.0 knots  True airspeed required at 50 feet 130 knots Results:  Ground speed at 50 feet 130 - 21.7 = 108.3 knots  Ground speed at touchdown 108.3 - 4.0 = 104.3 knots  True airspeed at touchdown 104.3 + 16.3 = 120.6 knots 3.1.9 Expansion of Takeoff and Landing Data for a Range of Airport Elevations. 3.1.9.1 These guidelines apply to expanding AFM takeoff and landing data above and below the altitude at which the airplane takeoff and landing performance tests are conducted. 3.1.9.2 Historically, limits were placed on the extrapolation of takeoff data. In the past, takeoff data could generally be extrapolated 6,000 feet above and 3,000 feet below the test field elevation when proven testing and data reduction methods were used. For extrapolations beyond these limits, a 2 percent takeoff distance penalty was to be applied for every additional 1,000 feet extrapolation. Such limitations were generally not applied to 3-11 ### Page 30 **AC 25-7D CHG 1** extrapolation of landing data, provided the effect of the higher true airspeed on landing distance was taken into account. 3.1.9.3 Since then, considerably more experience has since been gained both in terms of modeling airplane and propulsion system (i.e., turbine engines and propellers, where appropriate) performance and in verifying the accuracy of these models for determining high (and low) altitude takeoff and landing performance. This experience has shown that the soundness of the extrapolation is primarily a function of the accuracy of the propulsion system performance model and its integration with the airplane drag model. The basic aerodynamic characteristics of the airplane do not change significantly with altitude or ambient temperature, and any such effects are readily taken into account by standard airplane performance modeling practices. 3.1.9.4 As a result, with installed propulsion system performance characteristics that have been adequately defined and verified, airplane takeoff and landing performance data obtained at one field elevation may be extrapolated to higher and lower altitudes within the limits of the operating envelope without applying additional performance conservatisms. It should be noted, however, that extrapolation of the propulsion system data used in the determination and validation of propulsion system performance characteristics is typically limited to 3,000 feet above the highest altitude at which propulsion system parameters were evaluated for the pertinent power/thrust setting. (See paragraph 4.1 of this AC for more information on an acceptable means of establishing and verifying installed propulsion system performance characteristics.) 3.1.9.5 Note that certification testing for operation at airports that are above 8,000 feet should also include functional tests of the cabin pressurization system in accordance with paragraph 20.1.2.3 of this AC. Consideration should be given to any sensitivity to, or dependency upon airport altitude, such as: engine and auxiliary power unit (APU) starting, passenger oxygen, autopilot, autoland, autothrottle system power/thrust set/operation. 3.1.10 Tailwind Takeoff and Landing. 3.1.10.1 Tailwind Velocities of 10 Knots or Less. Approval may be given for performance, controllability, and engine operating characteristics for operations in reported tailwind velocities up to 10 knots without conducting additional flight tests at specific wind speeds. 3-12 ### Page 31 **AC 25-7D CHG 1** 3.1.10.2 Tailwind Velocities Greater than 10 Knots, up to 15 knots. 3.1.10.2.1 Performance. It is considered that takeoff, rejected takeoff, and landing distances, measured in tailwind conditions greater than 10 knots, are unreliable for use in determining airplane performance. Wind conditions of such magnitude are generally not sufficiently consistent over the length of the runway or over the time period required to perform the test maneuver. The 150 percent operational tailwind factor, required by §§ 25.105(d)(1) and 25.125(f), provides a satisfactory level of safety for operation in tailwinds up to 15 knots when using AFM data based on flight tests in nominally calm wind conditions. Tailwind approvals for greater than 15 knots should include dedicated takeoff and landing testing in elevated tailwind conditions to define AFM performance models. Note: The design requirements of § 25.479, Level landing conditions, also require the effects of increased contact speeds to be investigated if approval for landings with tailwinds greater than 10 knots is desired. 3.1.10.2.2 Control Characteristics. The test tailwind velocity for demonstrating handling qualities should be equal to the proposed limit tailwind increased by 5 knots. The intent of the 5-knot increase is to provide adequate margin for wind variability in operations, including currency of the wind data, averaging of the data by the measuring and reporting method, and the highly variable nature of higher wind conditions. Therefore, the test wind condition of 5 knots higher than the proposed tailwind limit should be an averaged or smoothed wind speed, not a peak wind speed. (Refer to Section 9.5.2.3 – Guidance to § 25.237 for explanation on averaged crosswind speed, to be adapted to tailwind.) Airplane control characteristics should be evaluated under the following conditions with the CG at the aft limit and the test mean tailwind velocity equal to the proposed limit tailwind increased by 5 knots: 1. Takeoff. Both all-engines-operating and one-engine-inoperative (i.e., with a simulated failure of the critical engine at the engine failure speed) takeoffs should be evaluated at a light weight with maximum approved takeoff flap deflection. 2. Landing. Approach and landing at light weight with maximum approved landing flap deflection. 3. Determination of the increased ground speed effect on gear vibration or shimmy, and flight director, or autopilot instrument landing system (ILS) approaches, terrain awareness warning system (TAWS) sink rate modes, etc. 4. Deviations above the glideslope should be shown as recoverable. In particular, it should be shown that the approach speed can be maintained, in tailwind conditions, on a glide path that is 1° steeper 3-13 ### Page 32 **AC 25-7D CHG 1** than a typical 3° glideslope. This can be shown by either flight test or analysis. Whatever method is used for glideslope recovery compliance, the actual tailwind (i.e., without correction to 10 meters height) need not be higher than the proposed tailwind limit increased by 5 knots. 3.1.10.2.3 Weight Limits. Consistent with the requirements of §§ 25.105(d)(1) and 25.125(f), the maximum takeoff and maximum quick turnaround weights should be determined using brake energies and tire speeds, as appropriate, calculated with the limit tailwind velocity factored by 150 percent. 3.1.10.2.4 Engine Operating Characteristics. Satisfactory engine operation should be demonstrated at the limit tailwind velocity factored by 150 percent. The demonstrations should include: 1. Zero groundspeed operation. 2. Takeoff power or thrust setting procedure used for AFM performance (typically completed by approximately 80 knots), both manually and automatically (autothrottle). 3. Reverse thrust operations. 3.1.10.2.5 Airplane Flight Manual. The AFM should contain a statement that the limitation for tailwinds greater than 10 knots reflects the capability of the airplane as evaluated in terms of airworthiness but does not constitute approval for operation in tailwinds exceeding 10 knots. 3.1.11 Procedures. 3.1.11.1 The performance-related flight test procedures are discussed in the following paragraphs of this AC:  Paragraph 4.2, Takeoff and Takeoff Speeds—§§ 25.105 and 25.107.  Paragraph 4.3, Accelerate-Stop Distance—§ 25.109.  Paragraph 4.4, Takeoff Path—§ 25.111.  Paragraph 4.5, Takeoff Distance and Takeoff Run—§ 25.113.  Paragraph 4.6, Takeoff Flight Path—§ 25.115.  Paragraph 4.7, Climb: General—§ 25.117.  Paragraph 4.8, Landing Climb: All-Engines-Operating—§ 25.119.  Paragraph 4.9, Climb: One-Engine-Inoperative—§ 25.121.  Paragraph 4.10, En Route Flight Paths—§ 25.123. 3-14 ### Page 33 **AC 25-7D CHG 1**  Paragraph 4.11, Landing—§ 25.125. 3.1.11.2 Performance Data for Multiple Flap or Additional Flap Positions. If approval of performance data is requested for flap settings at which no test data are available, the data may be obtained from interpolation of flight data obtained at no less than four flap settings that are within a constant configuration of other lift devices. If the span of flap settings is small and previously obtained data provide sufficient confidence (i.e., the shapes of the curves are known and lend themselves to accurate interpolation), data from three flap settings may be acceptable. 3.1.11.3 Flight Characteristics for Abnormal Configurations. See § 25.671(c). 3.1.11.3.1 For purposes of this AC, an abnormal configuration is an operational configuration that results from any single failure, or any combination of failures not shown to be improbable. 3.1.11.3.2 Flight characteristics for abnormal configurations may be determined by test or analysis to assure that the airplane is capable of continued safe flight and landing. Flight tests, if required, should be conducted at the critical conditions of altitude, weight, CG, and engine power or thrust associated with the configuration, and at the most critical airspeed between the speed reached one second after stall warning occurs (see paragraph 8.1.5.2.8 of this AC) and the maximum operating airspeed for the configuration. 3.2 Load Distribution Limits—§ 25.23. [Reserved] 3.3 Weight Limits and Center of Gravity Limits—§§ 25.25 and 25.27. [Reserved] 3.4 Empty Weight and Corresponding Center of Gravity—§ 25.29. [Reserved] 3.5 Removable Ballast—§ 25.31. 3.5.1 Explanation. None. 3.5.2 Procedures. Ballast may be carried during the flight tests whenever it is necessary to achieve a specific weight and CG location. Consideration should be given to the vertical as well as horizontal location of the ballast in cases where it may have an appreciable effect on 3-15 ### Page 34 **AC 25-7D CHG 1** the flying qualities of the airplane. The strength of the supporting structures should be considered in order to make sure they do not fail as a result of the anticipated loads that may be imposed during the particular tests. As required by § 21.35(a), applicants must show that these structures comply with the applicable structural requirements of part 25 before conducting flight tests with these structures in place. 3.6 Propeller Speed and Pitch Limits—§ 25.33. 3.6.1 Explanation. None. 3.6.2 Procedures. The tachometers and the airspeed indicating system of the test airplane should have been calibrated within the last six months. With that prerequisite satisfied, the following should be accomplished: 3.6.2.1 Determine that the propeller speeds and pitch settings are safe and satisfactory during all tests that are conducted in the flight test program within the certification limits of the airplane, engine, and propeller. This includes establishing acceptable low pitch (flight idle) blade angles on turbopropeller airplanes and verifying that propeller configurations are satisfactory at VMO/MMO to prevent propeller overspeed. 3.6.2.2 Determine that the propeller speeds and pitch settings are safe and satisfactory during all tests that are conducted to satisfy the performance requirements. 3.6.2.3 With the propeller governors operative and the propeller controls in full high revolutions per minute (RPM) position, determine that the maximum takeoff power settings do not exceed the rated takeoff RPM of each engine during takeoff and climb at the best rate-of-climb speed. 3.6.2.4 With the propeller governors made inoperative by mechanical means, determine the maximum power, no-wind, static RPMs. With the propeller governors operating on the low pitch stop, the engine speeds must not exceed 103 percent of the maximum allowable takeoff RPM or 99 percent of an approved maximum overspeed, as required by § 25.33(c). On turbopropeller engines, the engine speeds should not exceed the maximum engine speeds allowed by engine and propeller type designs. Note which systems were disabled and how the disablement was done. If maximum takeoff power torque or sea level standard conditions cannot be obtained on the test day, correct the data to these conditions by an acceptable means. A no-wind condition is considered to be a wind of 5 knots or less. The static RPM should be the average obtained with a direct crosswind from the left and a direct crosswind from the right. 3-16 ### Page 35 **AC 25-7D CHG 1** 3.6.2.5 If the above determinations are satisfactory, then measure the low-pitch stop setting and the high-pitch stop setting. These data may have been obtained from the propeller manufacturer and may be used, provided the pitch stops have not been changed since the manufacturer delivered the propeller. If measured, the blade station should be recorded. Include these blade angles in the type certificate data sheet. 3-17 ### Page 36 **CHAPTER** 4. FLIGHT: PERFORMANCE 4.1 General—§ 25.101. 4.1.1 Explanation of Propulsion System Behavior. Section 25.101(c) requires that airplane performance “correspond to the propulsive thrust available under the particular ambient atmospheric conditions, the particular flight conditions….” The propulsion systems (i.e., turbine engines and propellers, where appropriate) installed performance characteristics are primarily a function of engine power or thrust setting, airspeed, propeller efficiency (where applicable), altitude, and ambient temperature. Determine the effects of each of these variables to establish the thrust available for airplane performance calculations. 4.1.2 Procedures. 4.1.2.1 The intent is to develop a model of propulsion system performance that covers the approved flight envelope. Further, it should be shown that the combination of the propulsion system performance model and the airplane performance model is validated by the takeoff performance test data, climb performance tests, and tests used to determine airplane drag. Installed propulsion system performance characteristics may be established via the following tests and analyses:  Steady-state engine power (or thrust) setting versus power (or thrust) testing. See paragraph 4.1.2.2.  Lapse rate takeoff testing to characterize the behavior of power or thrust setting, rotor speeds, propeller effects (i.e., torque, RPM, and blade angle), or gas temperature as a function of time, thermal state, or airspeed, as appropriate. See paragraph 4.1.2.3.  Power/thrust calculation substantiation. See paragraph 4.1.2.4.  Effects of ambient temperature. See paragraph 4.1.2.5. 4.1.2.2 Steady-State Engine Power (or Thrust) Setting versus Power (or Thrust) Testing. Engines should be equipped with adequate instrumentation to allow the determination of thrust (or power). Data should be acquired to validate the model, including propeller-installed thrust, if applicable, over the range of power or thrust settings, altitudes, temperatures, and airspeeds for which approval is sought. Although it is not possible to definitively list or foresee all of the types of instrumentation that might be considered adequate for determining thrust (or power) output, two examples used in past certification programs are (1) engine pressure rakes, with engines calibrated in a ground test cell, and (2) fan speed, with engines calibrated in a ground test cell and the calibration data validated by the use of a flying test bed. In any case, the applicant should substantiate the adequacy 4-1 ### Page 37 of the instrumentation to be used for determining the thrust (or power) output. 4.1.2.3 Lapse Rate Takeoff Testing to Characterize the Behavior of Power or Thrust Setting, Rotor Speeds, Propeller Effects, or Gas Temperature as a Function of Time, Thermal State, or Airspeed. These tests should include the operation of an automatic takeoff thrust control system (ATTCS), if applicable, and should cover the range of power or thrust settings for which approval is sought. 4.1.2.3.1 Data for higher altitude power or thrust settings may be acquired via overboost (i.e., operating at a higher-than-normal power or thrust setting for the conditions) with the consent of the engine and propeller manufacturer(s), when applicable. When considering the use of overboost on turbopropeller propulsion system installations to stimulate higher altitude and ambient temperature range conditions, the capability to achieve an appropriate simulation should be evaluated based on the engine and propeller control system(s) and aircraft performance and structural considerations. Engine (gearbox) torque, rotor speed, or gas temperature limits, including protection devices to prohibit or limit exceedances, may prevent the required amount of overboost needed for performance at the maximum airport altitude sought for approval. Overboost may be considered as increased torque, reduced propeller speed, or a combination of both, in order to achieve the appropriate blade angle for the higher altitude and ambient temperature range simulation. Consideration for extrapolations will depend on the applicant’s substantiation of the proper turbopropeller propulsion system simulated test conditions. 4.1.2.3.2 Lapse rate characteristics should be validated by takeoff demonstrations at the maximum airport altitude for which takeoff approval is being sought. Alternatively, if overboost (see paragraph above) is used to simulate the power or thrust setting parameters of the maximum airport altitude for which takeoff approval is sought, the takeoff demonstrations of lapse rate characteristics can be performed at an airport altitude up to 3,000 feet lower than the maximum airport altitude. 4.1.2.4 Power/Thrust Calculation Substantiation. Installed power or thrust should be calculated via a mathematical model of the propulsion system, or other appropriate means, adjusted as necessary to match the measured inflight performance characteristics of the installed propulsion system. The propulsion system mathematical model should define the relationship of power or thrust to the power or thrust setting parameter over the range of power or thrust settings, airspeeds, altitudes, and temperatures for which approval is sought. For turbojet airplanes, the propulsion system mathematical model should be substantiated by ground tests in which thrust is directly measured via a calibrated load cell or equivalent means. For turbopropeller airplanes, the engine power 4-2 ### Page 38 measurement should be substantiated by a calibrated dynamometer or equivalent means, the engine jet thrust should be established by an acceptable engine model, and the propeller thrust and power characteristics should be substantiated by wind tunnel testing or equivalent means. 4.1.2.5 Effects of Ambient Temperature. The flight tests of paragraph 4.1.2.2 of this AC will typically provide data over a broad range of ambient temperatures. Additional data may be obtained from other flight or ground tests of the same type or series of engine. The objective is to confirm that the propulsion system model accurately reflects the effect of temperature over the range of ambient temperatures for which approval is being sought (operating envelope). Because thrust (or power) data can usually be normalized versus temperature using either dimensionless variables (e.g., theta exponents or a thermodynamic cycle model), it is usually unnecessary to obtain data over the entire ambient temperature range. There is no needed to conduct additional testing if:  The data show that the behavior of power or thrust and limiting parameters versus ambient temperature can be predicted accurately, and  Analysis based upon the test data shows that the propulsion system will operate at rated power or thrust without exceeding propulsion system limits. 4.1.2.6 Extrapolation of propulsion system performance data to 3,000 feet above the highest airport altitude tested (but no higher than the maximum takeoff airport altitude to be approved) is acceptable, provided the supporting data, including flight test and propulsion system operations data (e.g., engine and propeller control, limits exceedance, and surge protection devices scheduling), substantiates the proposed extrapolation procedures. Considerations for extrapolation depend upon an applicant’s determination, understanding, and substantiation of the critical operating modes of the propulsion system. This understanding includes a determination and quantification of the effects that propulsion system installation and variations in ambient conditions have on these modes. 4.2 Takeoff and Takeoff Speeds—§§ 25.105 and 25.107. 4.2.1 Explanation. Section 25.105 specifies the conditions that must be considered in determining the takeoff speeds, accelerate-stop distances, takeoff path, takeoff distance, and takeoff run in accordance with part 25 requirements. The primary objective of the takeoff tests required by § 25.107 is to determine the takeoff speeds for all takeoff configurations at 4-3 ### Page 39 **AC 25-7D CHG 1** all weight, altitude, and temperature conditions within the operational limits selected by the applicant. 4.2.2 Background Information. This information is not related to means of compliance for § 25.105 and § 25.107 but is included to increase awareness of the risks associated with in-ground-effect stall during execution of the required flight testing. The information is based upon previous industry experience and recommended best practices, much of which resulted from investigations following a flight test accident during takeoff performance testing. It is recommended that the following considerations be included in the applicant’s flight test hazard assessment in preparation for the takeoff performance flight testing. Ground Effect Considerations. It is important to understand the aerodynamic characteristics of the airplane in ground effect to avoid an inadvertent stall while operating at high angles of attack in close proximity to the ground. Ground effect is the phenomenon that modifies a body’s aerodynamic characteristics when it is generating lift close to the ground. The proximity of the ground suppresses downwash and wing tip vortices and affects the flow between the airplane and the ground, causing a slowing of the airstream under the wing, producing a positive pressure or buoyancy field. The impact of ground effect is largely dependent on the height of the wing above the ground and the magnitude of the resultant pressure field that exists between the underside of the wing and the ground plane. When the distance between the wing and ground is small (typically 5 to 10 percent of wingspan), the pressure field and extent of upwash/downwash suppression is large. When a significant blockage effect is experienced, the excess flow that cannot pass under the wing traverses the wing upper surface causing increased suction on the wing leading edge and increased lift at low angles of attack; this effect grows with larger and more powerful trailing edge flaps and their proximity to ground at liftoff attitude. The increased leading-edge suction tends to steepen the adverse pressure gradient behind the suction peak that can promote early aerodynamic flow separation, especially if the wing is already characterized by a leading-edge stall. It is especially important to note that when this effect is large, it will not only reduce the stall angle-of-attack, but it can also consequently decrease the maximum lift of the airplane in ground effect below the value observed in free air. It is also important to understand the effect that the Mach number has on the stall angle-of-attack in the range expected for takeoff. Experience has shown that reductions in the stall angle-of-attack with an increasing Mach number, observed in free air conditions, can also be considered applicable in ground effect. 4-4 ### Page 40 **AC 25-7D CHG 1** Due to remaining uncertainties in the in-ground-effect stall angle-ofattack, the test pilots involved in the takeoff testing should be familiar with the free air stall characteristics of the airplane, any predicted changes in the stall characteristics in ground-effect, and have established stall recovery techniques (e.g., reduction in angle-of-attack, advancing the throttles to increase thrust) in the event that a stall is encountered, including allowance for stall angle-of-attack hysteresis to re-attach the airflow in the recovery. 4.2.2.1 Ground Effect Estimation Methods: Industry experience has shown that a reduction in stall angle-of-attack in ground effect relative to free air of 4 to 5 degrees for takeoff flap deflections is a reasonable estimate in the absence of more configuration-specific data, though due to configuration differences there is no assurance that this estimate will be conservative in all cases. Wind tunnel results obtained at lower Reynolds numbers than full-scale flight should be used with extreme caution when predicting ground effect. While the “linear range” effects of such sub-scale testing can be accurate, the impact of ground effect on stall may not be captured at low Reynolds numbers. Computational fluid dynamics (CFD) results have been shown to produce useful indications of in-ground-effect stall, but the undertaking of predicting stall angles with CFD is, in general, not trivial. Before using wind tunnel or CFD predicted in-ground-effect stall angles, comparisons should be made with results from free air stall flight testing to understand and/or improve the accuracy of the predictions. Caution should be exercised to ensure that conservative margins are identified for flight test use. 4.2.3 Procedures: General. 4.2.3.1 Section 25.105(c)(1) requires the takeoff performance data to be determined for smooth, dry and wet, hard-surfaced runways. Paragraph 4.3 of this AC describes methods for determining the accelerate-stop distances required by § 25.109. Paragraph 4.5 describes methods for determining the takeoff distance and takeoff run required by § 25.113. 4.2.3.2 In accordance with § 25.101(f), testing for determining the accelerate-stop distances, takeoff flight paths, and takeoff distances should be accomplished using procedures established by the applicant for operation in service. In accordance with §25.101(h), these procedures must be able to be consistently executed in service by crews of average skill, use methods or devices that are safe and reliable, and include allowances for any time delays in the execution of the procedures that may reasonably be expected in service. These requirements prohibit the use of exceptional piloting techniques, such as higher control force inputs or higher pitch 4-5 ### Page 41 **AC 25-7D CHG 1** rates than would occur in operational service, from being used to generate unrealistic takeoff distances. The intent of these requirements is to establish takeoff performance representative of that which can reasonably be expected to be achieved in operational service. 4.2.3.3 Attention should be paid to all potential sources of airspeed error, but special consideration should be given to airplanes with electronic instruments in the cockpit that apply electronic filtering to the airspeed data. This filtering, which causes a time delay in the airspeed indication, can be a source of significant systematic error in the presentation of airspeed to the flightcrew. With normal takeoff acceleration, the airplane will be at a higher speed than is indicated by the cockpit instrument, which can result in longer distances than are presented in the AFM, particularly in the event of a rejected takeoff near the indicated V1 speed. The effects of any time delays caused by electronic filtering, pneumatic system lag, or other sources should be adequately addressed in the AFM speed and distance presentations. Further explanation of airspeed lag, particularly pertaining to airplanes with electronic instruments in the cockpit, and procedures for calibrating the airspeed indicating system (§ 25.1323(b)) are presented in paragraph 33.3 of this AC. 4.2.4 Procedures: Section 25.107(a)(1)—Engine Failure Speed (VEF). The engine failure speed (VEF) is defined as the calibrated airspeed at which the critical engine is assumed to fail and must be selected by the applicant. VEF cannot be less than the ground minimum control speed (VMCG). 4.2.5 Procedures: Section 25.107(a)(2)—V1. V1 may not be less than VEF plus the speed gained with the critical engine inoperative during the time interval between VEF and the instant at which the pilot takes action after recognizing the engine failure. This is indicated by pilot application of the first deceleration means such as brakes, throttles, spoilers, etc. during accelerate-stop tests. The applicant may choose the sequence of events. Refer to paragraph 4.3 of this AC, addressing § 25.109, for a more complete description of rejected takeoff (RTO) transition procedures and associated time delays. 4.2.6 Procedures: Section 25.107(b)—Minimum Takeoff Safety Speed (V2MIN). 4.2.6.1 V2MIN, in terms of calibrated airspeed, cannot be less than: 4.2.6.1.1 1.1 times the VMC defined in § 25.149. 4.2.6.1.2 1.13 times VSR for two-engine and three-engine turbopropeller and reciprocating engine-powered airplanes and for all turbojet airplanes that do not have provisions for obtaining a significant reduction in the one-engine-inoperative power-on stalling speed (i.e., boundary layer 4-6 ### Page 42 **AC 25-7D CHG 1** control, blown flaps, etc.). The value of VSR to be used in determining V2MIN is the free air reference stall speed in the applicable takeoff configuration, landing gear retracted, except for those airplanes with a fixed landing gear or for gear-down dispatch. 4.2.6.2 V2MIN may be reduced to 1.08 times VSR for turbopropeller and reciprocating engine-powered airplanes with more than three engines, and turbojet powered airplanes with adequate provisions for obtaining significant power-on reference stall speed reduction through the use of such things as boundary layer control, blown flaps, etc. 4.2.6.3 For propeller-driven airplanes, the difference between the two margins, based upon the number of engines installed on the airplane, is because the application of power ordinarily reduces the stalling speed appreciably. In the case of the two-engine propeller-driven airplane, at least half of this reduction is eliminated by the failure of an engine. The difference in the required factors therefore provides approximately the same margin over the actual stalling speed under the power-on conditions that are obtained after the loss of an engine, no matter what the number of engines (in excess of one) may be. Unlike the propeller-driven airplane, the turbojet/turbofan powered airplane does not show any appreciable difference between the power-on and power-off stalling speed. This is due to the absence of the propeller, which ordinarily induces a slipstream with the application of power causing the wing to retain its lift to a speed lower than the power-off stalling speed. The applicant’s selection of the two speeds specified will influence the nature of the testing required in establishing the takeoff flight path. 4.2.7 Procedures: Section 25.107(c)—Takeoff Safety Speed (V2). 4.2.7.1 V2 is the calibrated airspeed that is attained at or before the airplane reaches a height of 35 feet above the takeoff surface after an engine failure at VEF using an established rotation speed (VR). From the liftoff point, the takeoff surface extends to the end of the takeoff distance continuing at the same slope as the runway. During the takeoff speeds demonstration, V2 should be continued to an altitude sufficient to assure stable conditions beyond the 35-foot height. V2 cannot be less than V2MIN. In addition, V2 cannot be less than the liftoff speed, VLOF, which is defined in § 25.107(f). In accordance with § 25.107(c), V2, in terms of calibrated airspeed, may not be less than VR plus the speed increment attained before reaching a height of 35 feet above the takeoff surface using a takeoff maneuver that can be executed consistently by crews of average skill per the requirement of § 25.101(h)(1), and a speed that provides the maneuvering capability specified in § 25.143(h). Due to the constraints on VR and VLOF specified in § 25.107(e), and accounting for other constraints required for safe 4-7 ### Page 43 **AC 25-7D CHG 1** operation, such as maintaining adequate margin to the in-ground-effect stall angle-of-attack during a dynamic takeoff maneuver, V2 may be forced to be greater than V2MIN for some, if not all, of the thrust/weight range of operation. In addition, § 25.111(c)(2) stipulates that the airplane must reach V2 before it is 35 feet above the takeoff surface and continue at a speed not less than V2 until it is 400 feet above the takeoff surface. These requirements were first expressed in Special Civil Air Regulation No. SR-422, Turbine-Powered Transport Category Airplanes of Current Design (SR-422A), paragraphs 4T.114(b)(4) and (c)(3) and 4T.116(e). The concern that the regulation change was addressing was the overshoot of V2 after liftoff under the previous requirement that the airplane attain V2 on, or near, the ground. The intent of the current requirement is to allow an acceleration to V2 after liftoff but not to allow a decrease in the field length required to attain a height of 35 feet above the takeoff surface by attaining a speed greater than V2, under low drag ground conditions, and using the excess kinetic energy to attain the 35-foot height. 4.2.7.2 In the case of turbojet powered airplanes, when most of the one-engine-inoperative data have been collected using throttle chops, V2, and its relationship to VR, should be substantiated by at least a limited number of fuel cuts at VEF. For derivative programs not involving a modification that would affect thrust decay characteristics, demonstrations of fuel cuts may be unnecessary. 4.2.7.3 For propeller-driven airplanes, the use of fuel cuts can be more important to ensure that the takeoff speeds and distances are obtained with the critical engine’s propeller attaining the position it would during a sudden engine failure. The number of tests that should be conducted using fuel cuts depends on the correlation obtained with the throttle chop data and substantiation that the data analysis methodology adequately models the effects of a sudden engine failure. 4.2.8 Procedures: Section 25.107(d)—Minimum Unstick Speed (VMU). 4.2.8.1 Section 25.107(d) states, “VMU speeds must be selected by the applicant.” An applicant can either determine the lowest possible VMU speeds or select a higher speed that supports the takeoff performance targets of the airplane. Regardless of how the applicant selects the VMU speeds, compliance must be shown with § 25.107(d), (e)(1)(iv), (e)(3), and (e)(4) to show that the selected VMU speeds allow the airplane to safely lift off the ground and continue the takeoff. 4.2.8.2 An applicant should comply with § 25.107(d) by conducting VMU tests with all engines operating and also with one engine inoperative. During these tests, the takeoff should be continued until the airplane is out of 4-8 ### Page 44 AC 25-7D CHG 1 ground effect. The airplane pitch attitude should not be decreased after liftoff. 4.2.8.3 VMU testing to demonstrate the lowest VMU speed is a maximum performance flight test maneuver, and liftoff may occur very near the angle-of-attack for maximum lift coefficient in ground effect, with the minimum margin to in-ground-effect stall occurring in the vicinity of liftoff. As discussed in paragraph 4.2.2 of this AC, to ensure flight test safety, a thorough understanding of the stall angle-of-attack in ground effect and appropriate angle-of-attack margins should be established and maintained during testing (see Figure 4-A). In addition, even though pitch attitude may be held fairly constant during the maneuver, environmental conditions and transiting through ground effect may result in changes in angle-of-attack. It is permissible to lift off at a speed that is below the normal stall warning speed, provided no more than light buffet is encountered. The use of a flight test device that restricts the on-ground pitch attitude has been found useful by some manufacturers in reducing the risk of over-rotation and in-ground-effect stall. 4.2.8.3.1 It is important for the flight test team to understand the control laws and any transitions between control laws during takeoff (e.g., based on weight on wheels) for an electronic flight control system that may present unique hazards that should be taken into account. 4.2.8.3.2 An artificial stall warning system (e.g., a stick shaker) may be disabled or adjusted to a more suitable value during VMU testing. 4.2.8.3.3 If the airplane is equipped with a stick pusher, high angle-of-attack limiting function, or other system that may affect the conduct of the test, the angle of attack setting for activation of the system may be selected by the applicant and differ from the nominal setting. The system may alternatively be disabled or its activation delayed for test purposes until a safe altitude is reached. However, for airplanes equipped with a stick pusher that is not designed to be inhibited during takeoff, the VMU test demonstrations will need to be assessed and will only remain valid if the stick pusher would not have activated with the angle-of-attack indication means set at the lowest angle within production tolerances. 4.2.8.3.4 Note that if there are changes to an airplane’s aerodynamic stall angle due to ground effect, then a stall warning system, stick pusher, or high angleof-attack limiting function that is not specifically designed to account for ground effect may not activate prior to an aerodynamic stall, during a VMU test. The airplane’s stall angle-of-attack and angle-of-attack sensor corrections, both in and out of ground effect, must be thoroughly understood by the test crew to determine if these systems will provide the benefit of protecting the airplane during VMU testing. 4-9 ### Page 45 AC 25-7D CHG 1 Figure 4-A. In-Ground-Effect Stall Margin 4.2.8.4 In lieu of conducting one-engine-inoperative VMU tests, the applicant may conduct all-engines-operating VMU tests if all pertinent factors that would be associated with an actual one-engine-inoperative VMU test are simulated or otherwise taken into account. To take into account all pertinent factors, it may be necessary to adjust the resulting VMU test values analytically. The factors to be accounted for should include at least the following:  Thrust/weight ratio for the one-engine-inoperative range.  Controllability (may be related to one-engine-inoperative free air tests, such as minimum control speed in the air (VMCA)).  Increased drag due to use of lateral/directional control systems.  Reduced lift due to use of devices such as wing spoilers for lateral control.  Adverse effects of use of any other systems or devices on control, drag, or lift. 4.2.8.5 The number of VMU tests needed may be minimized by testing only the critical all-engines-operating and one-engine-inoperative thrust/weight ratios, provided the VMU speeds determined at these critical conditions are used for the range of thrust/weights appropriate to the all-engines-operating and one-engine-inoperative configurations. The critical thrust/weight is established by correcting, to the VMU speed, the thrust that results in the airplane achieving its limiting one-engine-inoperative climb gradient at the normally scheduled speed and in the appropriate configuration. 4-10 ### Page 46 **AC 25-7D CHG 1** 4.2.8.6 Amendment 25-42, effective March 1, 1978, revised §§ 25.107(d) and 25.107(e)(1)(iv) in order to permit the one-engine-inoperative VMU to be determined by all-engines-operating tests at the thrust/weight ratio corresponding to the one-engine-inoperative condition. As revised, § 25.107(d) specifies that VMU must be selected for the range of thrust/weight ratios to be certificated, rather than for the all-enginesoperating and one-engine-inoperative conditions as was previously required. In determining the all-engines-operating thrust/weight ratio that corresponds to the one-engine-inoperative condition, consider trim and control drag differences between the two configurations in addition to the effect of the number of engines operating. The minimum thrust/weight ratio to be certificated is established by correcting, to the VMU speed, the thrust that results in the airplane achieving its limiting engine-out climb gradient in the appropriate configuration and at the normally scheduled speed. 4.2.8.7 To conduct the VMU tests, rotate the airplane as necessary to achieve the VMU attitude. It is acceptable to use some additional nose-up trim over the normal trim setting during VMU demonstrations. Even on airplanes that have sufficient control authority to achieve the target VMU attitude prior to liftoff, the use of additional nose-up trim can be beneficial to the conduct of the test by giving the pilot additional time to stabilize on the VMU attitude prior to lifting off. If additional nose-up trim is used, the additional considerations of paragraph 4.2.8.8 below apply. VMU is the speed at which the weight of the airplane is completely supported by aerodynamic lift and thrust forces. Some judgment may be necessary on airplanes that have tilting main landing gear bogies. Determining the liftoff point from gear loads and wheel speeds has been found acceptable in past programs. After liftoff, the airplane should be flown out of ground effect. During liftoff and the subsequent climbout, the airplane should be fully controllable. 4.2.8.8 VMU Testing for Airplanes Having Limited Pitch Control Authority. 4.2.8.8.1 For some airplanes with limited pitch control authority, it may not be possible, at forward CG and normal trim, to rotate the airplane to a liftoff attitude where the airplane could otherwise perform a clean flyaway at a minimum speed had the required attitude been achieved. This may occur only over a portion of the takeoff weight range in some configurations. Though generally associated with the inability of the pitch control surfaces to provide adequate pitching moment to rotate the airplane to the desired pitch attitude at low thrust/weight ratio conditions, the same phenomenon may occur at high thrust/weight ratio conditions for airplanes with high thrust lines (e.g., aft engines mounted high on the fuselage). When limited pitch control authority is clearly shown to be the case, VMU test conditions 4-11 ### Page 47 AC 25-7D CHG 1 may be modified to allow testing aft of the forward CG limit and/or with use of more airplane nose-up trim than normal. The VMU data determined with this procedure should be corrected to those values representative of the appropriate forward limit; the variation of VMU with CG may be assumed to be like the variation of free air stalling speed with CG. Although the development of scheduled takeoff speeds may proceed from these corrected VMU data, additional tests are required (see paragraph 4.2.8.8.2 below) to check that the relaxed VMU criteria have not neglected problems that might arise from operational variations in rotating airplanes with limited pitch control authority. 4.2.8.8.2 In the following assurance test, the airplane should demonstrate safe flyaway characteristics: Minimum speed liftoff should be demonstrated at the critical forward CG limit with normal trim. For airplanes with a cutback forward CG at heavy weight, two weight/CG conditions should be considered. The heavy weight tests should be conducted at maximum structural or maximum sea level climb-limited weight with the associated forward CG. The full forward CG tests should be conducted at the highest associated weight. Alternatively, testing may be conducted at a single weight if an analysis is provided that identifies the critical weight/CG combination with regard to limited pitch attitude capability for liftoff. These assurance tests should be conducted at the thrust/weight ratio that is most critical for attaining a pitch attitude that will provide a minimum liftoff speed. 4.2.8.8.3 For airplanes that are limited by low thrust/weight conditions, tests should be conducted at the minimum thrust/weight ratio for both the simulated one-engine-inoperative test (i.e., symmetrical reduced thrust) case and the all-engines-operating case. 4.2.8.8.4 For airplanes that are limited by high thrust/weight conditions, tests should be conducted at the highest thrust/weight ratio within the airplane’s operating envelope for both the simulated one-engine-inoperative case (i.e., symmetrical reduced thrust) and the all-engines-operating case. 4.2.8.8.5 One acceptable test technique is to hold full nose-up control column as the airplane accelerates. As pitch attitude is achieved to establish the minimum liftoff speed, pitch control may be adjusted to prevent over-rotation, but the liftoff attitude should be maintained as the airplane flies off the ground and out of ground effect. 4.2.8.8.6 The resulting liftoff speeds are acceptable if the test proves successful and the liftoff speed is at least 5 knots below the normally scheduled liftoff speed. 4-12 ### Page 48 **AC 25-7D CHG 1** 4.2.8.8.7 This minimum 5 knot margin from the scheduled liftoff speed provides some leeway for operational variations such as mistrim, CG errors, etc., that could further limit the elevator authority. The reduced VMU margins arising from this test, relative to those specified in § 25.107(e)(1)(iv), are considered acceptable because of the reduced probability of a pitch control authority-limited airplane getting into a high drag condition due to over-rotation. 4.2.8.9 VMU Testing for Geometry Limited Airplanes. 4.2.8.9.1 For airplanes that are geometry limited (i.e., the minimum possible VMU speeds are limited by tail contact with the runway), § 25.107(e)(1)(iv)(B) allows the VMU to VLOF speed margins to be reduced to 108 percent and 104 percent for the all-engines-operating and one-engine-inoperative conditions, respectively. The VMU demonstrated should be sound and repeatable. As discussed in paragraph 4.2.2 of this AC, to ensure flight test safety, a thorough understanding of the estimated stall angle-of-attack in ground effect and appropriate angle-of-attack margins should be established and maintained during testing. 4.2.8.9.2 An airplane that is deemed to be geometry limited at the conditions tested is expected to be geometry limited over its entire takeoff operating envelope. If this is not the case, the airplane is not considered geometry limited and the reduced VMU to VLOF speed margins do not apply. Also, if a flight-test device is used with the intent to artificially restrict the rotation attitude (typically more than 0.5 degrees below the production configuration) to prevent over-rotation, and the airplane would not otherwise be geometry limited, this airplane would not be considered geometry limited and the reduced VMU to VLOF speed margins do not apply. 4.2.8.9.3 One acceptable means for demonstrating compliance with §§ 25.107(d) and 25.107(e)(1)(iv) with respect to the capability for a safe liftoff and fly-away from the geometry limited condition is to show that at the lowest thrust-to-weight ratio for the all-engines-operating condition: 1. In the speed range from 96 to 100 percent of the actual liftoff speed, the aft under-surface of the airplane should be in contact with the runway. Because of the dynamic nature of the test, it is recognized that contact will probably not be maintained during this entire speed range, so some judgment is necessary. It has been found acceptable for contact to exist approximately 50 percent of the time that the airplane is in this speed range. 2. Beyond the point of liftoff to a height of 35 feet, the airplane’s pitch attitude should not decrease below that at the point of liftoff, nor should the speed increase more than 10 percent. 4-13 ### Page 49 **AC 25-7D CHG 1** 3. The horizontal distance from the start of the takeoff to a height of 35 feet above the takeoff surface should not be greater than 105 percent of the distance determined in accordance with § 25.113(a)(2) without applying the 115 percent factor. 4.2.8.10 VMU for a Stretched Version of a Tested Airplane. 4.2.8.10.1 VMU speeds obtained by flight testing one model of an airplane type may be used to generate VMU speeds for a geometry-limited stretched version of that airplane. If the short body airplane met the criteria for the 104/108 percent VMU/VLOF speed margin for geometry limited airplanes as permitted by § 25.107(e)(1)(iv)(B) and discussed in paragraph 4.2.8.9.1 above, then the flight tests described in paragraph 4.2.8.10.2 below should be performed on the stretched derivative. Otherwise, the flight tests described in paragraph 4.2.8.9.3 above should be performed on the stretched derivative. 4.2.8.10.2 Since the concern for tail strikes is increased with the stretched airplane, the following should be accomplished, in addition to normal takeoff tests, when the VMU schedule of the stretched derivative is derived from that of the shorter body parent airplane: 1. The VMU of the stretched derivative airplane should be determined by correcting the VMU of the shorter body tested airplane for the reduced runway pitch attitude capability and revised CG range of the stretched airplane. Alternatively, stretched airplane VMU speeds not determined in this manner should be substantiated by flight testing or a rational analysis. Scheduled rotation speeds for the stretched airplane should result in at least the required liftoff speed margins above the corrected VMU required by § 25.107(e)(1)(iv) for the one-engine-inoperative and all-engines-operating takeoff conditions. 2. At both the forward and aft CG limits, and over the thrust-to-weight range for each takeoff flap, the following takeoff tests should be accomplished. The tests described in paragraphs a and b below should be accomplished with not more than occasional, minor (i.e., non-damaging) tail strikes. As discussed in paragraph 4.2.2 of this AC, to ensure flight test safety, a thorough understanding of the estimated stall angle-of-attack in ground effect and appropriate angle-of-attack margins should be established and maintained during testing. a. All-engines-operating, early rotation tests specified in paragraph 4.2.9.3.2, including both the rapid rotations and over-rotations as separate test conditions. 4-14 ### Page 50 **AC 25-7D CHG 1** b. One-engine-inoperative, early rotation tests specified in paragraph 4.2.9.2 below. c. All-engines-operating, moderate rotation rate (i.e., more rapid than normal) takeoff tests, using the scheduled VR and normal pitch attitude after liftoff. Tail strikes should not occur for this condition. 4.2.9 Procedures: Section 25.107(e)—Rotation Speed (VR). 4.2.9.1 VR in terms of calibrated airspeed, must be selected by the applicant. VR has a number of constraints that must be observed in order to comply with § 25.107(e): 4.2.9.1.1 VR may not be less than V1; however, it can be equal to V1 in some cases. 4.2.9.1.2 VR may not be less than 105 percent of VMCA. 4.2.9.1.3 VR must be a speed that will allow the airplane to reach V2 at or before reaching a height of 35 feet above the takeoff surface, when the takeoff is conducted using normal takeoff procedures. 4.2.9.1.4 VR must be a speed that will result in liftoff at a speed not less than 110 percent of VMU (unless geometry limited) for the all-engines-operating condition, and not less than 105 percent of the VMU (unless geometry limited) determined at the thrust/weight ratio corresponding to the one-engine-inoperative condition for each set of conditions, such as weight, altitude, temperature, and configuration when the airplane is rotated at its maximum practicable rate. For this requirement, maximum practicable rate depends on the airplane configuration, type of pitch controller, flight control system design, and the takeoff procedure. The rotation rate need not be increased beyond the point that prevents capturing the normal takeoff rotation attitude, without using exceptional piloting skill or strength. Rotation rates between 120% and 150% of the nominal rate used in determination of the takeoff performance in accordance with §§ 25.105, 25.111, and 25.113 have previously been found acceptable. Alternatively, this rotation rate can be determined analytically as a representatively high rotation rate from a significant sampling of takeoffs performed during the flight test program, including the takeoff field-performance tests. 4.2.9.1.5 VR may not be less than the speed necessary to demonstrate the oneengine inoperative and all-engines-operating in-service variation tests (early rotation, over-rotation, out-of-trim) from the requirements of § 25.107(e)(3) and (4), which specify that expected variations in service should not result in any unsafe flight characteristics or a marked increase 4-15 ### Page 51 AC 25-7D CHG 1 in takeoff distance. As discussed in paragraph 4.2.2 of this AC, to ensure flight test safety, a thorough understanding of the estimated stall angle-ofattack in ground effect and appropriate angle-of-attack margins should be established and maintained during testing. It should be noted that ensuring successful demonstrations for these in-service variation criteria may in some cases require increasing VR (and thus V2) to a higher speed than what would otherwise be required by § 25.107(e)(1). 4.2.9.2 Early Rotation, One-Engine-Inoperative Test. 4.2.9.2.1 In showing compliance with § 25.107(e)(3), some guidance relative to the airspeed attained at the 35-foot height during the associated flight test is necessary. As this requirement only specifies an early rotation (VR-5 knots), it is interpreted that pilot technique is to remain the same as normally used for a one-engine-inoperative condition. With these considerations in mind, it is apparent that the airspeed achieved at the 35-foot point can be somewhat below the normal scheduled V2 speed. However, the amount of permissible V2 speed reduction should be limited to a reasonable amount as described below. 4.2.9.2.2 These test criteria apply to all unapproved, new, basic model airplanes. They also apply to previously approved airplanes when subsequent testing is warranted. However, for those airplanes where these criteria are more stringent than those previously applied, consideration will be given to permitting some latitude in the test criteria. 4.2.9.2.3 In conducting the flight tests required by § 25.107(e)(3), the test pilot should use the normal/natural rotation technique associated with the use of scheduled takeoff speeds for the airplane being tested. Intentional tail or tail skid contact is not considered acceptable. Non-damaging contact due to inadvertent over-rotation is acceptable provided there is a prompt recovery to the normal one-engine-inoperative takeoff pitch attitude. Further, the airspeed attained at the 35-foot height during this test should not be less than the scheduled V2 value minus 5 knots. These speed limits should not be considered or used as target V2 test speeds, but rather are intended to provide an acceptable range of speed departure below the scheduled V2 value. To ensure flight test safety, the maximum angle-ofattack, as a function of height above ground expected for this maneuver, should be confirmed to either fall below those previously demonstrated by VMU tests, or provide conservative margin to predicted in-ground-effect stall angle-of-attack as a function of height above ground (see Figure 4-A). Note: experience has shown that the lowest margin to in-ground-effect stall angle-of-attack occurs in the vicinity of liftoff and that asymmetric flight conditions may adversely affect the stall angle-of-attack. 4-16 ### Page 52 **AC 25-7D CHG 1** 4.2.9.2.4 In this test, the simulated engine failure should be accomplished sufficiently in advance of the VR test speed to allow for engine spin-down, unless this would be below the VMCG, in which case VMCG should govern. The normal one-engine-inoperative takeoff distance may be analytically adjusted to compensate for the effect of the early power or thrust reduction. Further, in those tests where the airspeed achieved at the 35-foot height is slightly less than the V2-5 knots limiting value, it will be permissible, in lieu of conducting the tests again, to analytically adjust the test distance to account for the excessive speed decrement. 4.2.9.3 All-Engines-Operating Tests. 4.2.9.3.1 Section 25.107(e)(4) states that there must not be a “marked increase” in the scheduled takeoff distance when reasonably expected service variations such as early and excessive rotation and out-of-trim conditions are encountered. This has been interpreted as requiring takeoff tests with all engines operating with:  A lower than scheduled rotation speed, and  Out-of-trim conditions, but with rotation at the scheduled VR speed. Note: The expression “marked increase” in the takeoff distance is considered to be any amount in excess of 1 percent of the scheduled takeoff distance. Thus, the tests should not result in field lengths more than 101 percent of the takeoff field lengths calculated in accordance with the applicable requirements of part 25 for presentation in the AFM. 4-17 ### Page 53 **AC 25-7D CHG 1** 4.2.9.3.2 For the early rotation condition with all engines operating, and at a weight as near as practicable to the maximum sea level standard day takeoff weight limit, it should be shown by tests that when the airplane is rotated at a speed below the scheduled VR, no “marked increase” in the scheduled AFM field length will result. For these tests, the airplane should be rotated at a speed equal to the scheduled VR minus 7 percent or the scheduled VR minus 10 knots, whichever results in the higher rotation speed. Tests should be conducted at (1) a rapid rotation rate to the normal takeoff attitude, and, as a separate set of tests, and (2) an over-rotation of 2° above the normal attitude following liftoff at a normal rotation rate. For this requirement, the rapid rotation rate achievable at VR-10 kt (or -7%) depends on the airplane configuration, type of pitch controller, flight control system design and the normal takeoff procedure. The rotation rate need not be increased beyond the point that prevents capturing the normal takeoff rotation attitude without using exceptional piloting skill or strength. For tests using over-rotations, the resulting increased pitch attitude should be maintained until the airplane is out of ground effect. Tail strikes during this demonstration are acceptable if they are minor and do not result in unsafe conditions. The maximum angle-of-attack as a function of height above ground expected for both of these maneuvers should be confirmed to either fall below those previously demonstrated by flight tests or provide conservative margin to the predicted in-groundeffect stall angle-of-attack as a function of height above ground (see Figure 4-A). Note: Experience has shown that the lowest margin to in-ground-effect stall angle-of-attack occurs in the vicinity of liftoff. 4.2.9.3.3 For reasonably expected out-of-trim conditions with all engines operating and as near as practicable to the maximum weight allowed under sea level standard day conditions, it should be shown that there will not be a “marked increase” in the scheduled AFM takeoff distance when rotation is initiated in a normal manner at the scheduled VR speed. The amount of mistrim should be the maximum mistrim that would not result in a takeoff configuration warning, including taking into account the takeoff configuration warning system rigging tolerance. It is permissible to accept an analysis in lieu of actual testing if the analysis shows that the out-of-trim condition would not present unsafe flight characteristics or a “marked increase” in the scheduled AFM field lengths. 4-18 ### Page 54 **AC 25-7D CHG 1** 4.2.9.3.4 Section 25.107(e)(4) also states that the reasonably expected variations in service from the established takeoff procedures for the operation of the airplane may not result in unsafe flight characteristics. For example, for an airplane loaded to obtain a forward CG position and mistrimmed for an aft CG loading, it may not be possible to rotate at the normal operating speeds due to excessive control force or lack of primary pitch control authority. This may result in an excessive delay in accomplishing the rotation. Such a condition would be considered an unsafe flight characteristic. Similarly, for an airplane loaded to obtain an aft CG position and mistrimmed for a forward CG loading, it may not be possible to readily arrest a self-rotating tendency. This rotation, if abrupt enough and rapid enough, could lead to stall. Qualitative assessments should be made by the test pilot in the following takeoff tests with all engines operating: 1. The test pilot should determine that no unsafe characteristics exist with the airplane loaded to the forward CG limit and the stabilizer mistrimmed in the airplane nose-down direction. The amount of mistrim should be the maximum mistrim that would not result in a configuration warning (including taking into account takeoff warning system tolerances). Rotation should be initiated at the scheduled rotation speed for the airplane weight and ambient conditions. Unsafe characteristics include an excessive pitch control force to obtain normal airplane response or an excessive time to achieve perceptible rotation. 2. The test pilot should determine that no unsafe characteristics exist with the airplane loaded to the aft CG limit and the stabilizer mistrimmed in the airplane nose-up direction. The amount of mistrim should be the maximum mistrim that would not result in a configuration warning (including taking into account takeoff warning system tolerances). The airplane should be rotated at the scheduled rotation speed for the airplane weight and ambient conditions. Unsafe characteristics include an abrupt self-rotating tendency that cannot be checked with normal control input, or an excessive pitch control force required to maintain the airplane in the normal pitch attitude prior to the scheduled rotation speed or during rotation and initial climb. 3. For the tests described in paragraphs 1 and 2 above, the flight characteristics should be assessed at the most critical combinations of airplane weight, wing flap position and engine power or thrust for the out of trim position being considered. 4-19 ### Page 55 **AC 25-7D CHG 1** 4.2.9.4 Stall Warning during Takeoff Speed Tests. The presumption is that if an operational pilot was to make an error in takeoff speeds that resulted in an encounter with stall warning, the likely response would be to recover aggressively to a safe flight condition rather than trying to duplicate the AFM takeoff flight path. Therefore, the activation of any stall warning devices, or the occurrence of airframe buffeting during takeoff speed testing, is unacceptable. 4.2.9.5 Stick Forces during Takeoff Speed Tests. According to § 25.143(a)(1) and (b), stick forces to initiate rotation and continue the takeoff during takeoff flight testing must comply with the control force limits of § 25.143(d). This includes the mistrim takeoff tests described in paragraphs 1 and 2 (on page 4-19) to show compliance with § 25.107(e)(4), which are considered to represent probable operating conditions under § 25.143(b). Stick forces should be those that result from using the takeoff procedures established by the manufacturer for use in operational service in accordance with § 25.101(f) and must comply with § 25.101(h). 4.2.10 Procedures: Section 25.107(f)—Liftoff Speed (VLOF). 4.2.10.1 VLOF is defined as the calibrated airspeed at which the airplane first becomes airborne (i.e., no contact with the runway). This allows comparison of liftoff speed with tire limit speed. VLOF differs from VMU in that VMU is the minimum possible VLOF speed for a given configuration, and depending upon landing gear design, VMU liftoff is shown to be the point where all of the airplane weight is being supported by airplane lift and thrust forces and not any portion by the landing gear. For example, after the VMU speed is reached, a truck tilt actuator may force a front or rear wheel set to be in contact with the runway, even though the liftoff is in progress by virtue of lift being greater than weight. 4.2.10.2 The maximum ground speed at liftoff, considering the entire takeoff operating envelope and taking into account 50 percent of the headwind and 150 percent of the tailwind, in accordance with § 25.105(d)(1), must not exceed the tire speed rating established under § 25.733(a) or (c). 4.3 Accelerate-Stop Distance—§ 25.109. 4.3.1 Explanation. 4.3.1.1 The accelerate-stop distance is the horizontal distance from a reference point on the airplane at initial brake release to that same reference point after the airplane is brought to a stop. 4-20 ### Page 56 **AC 25-7D CHG 1** 4.3.1.2 This section describes test demonstrations and data expansion methods necessary to determine accelerate-stop distances for publication in the FAA-approved AFM, as required by § 25.1583(h) (by reference to § 25.1533). Amendment 25-92 revised some aspects of the part 25 accelerate-stop criteria and added new requirements related to the stopping capability of the airplane as affected by brake wear and wet runways. The changes imparted to the accelerate-stop requirements by amendment 25-92 are listed below. (For other material related to the use of accelerate-stop distances, see 14 CFR parts 121 and 135.) 4.3.1.2.1 Section 25.101(i) was added to require accelerate-stop distances to be determined with all the airplane wheel brake assemblies at the fully worn limit of their allowable wear range. 4.3.1.2.2 Section 25.105(c)(1) was revised to require takeoff data to be determined for wet, in addition to dry, hard surfaced runways. At the applicant’s option, takeoff data may also be determined for wet runways that have grooved or porous friction course surfaces. 4.3.1.2.3 Section 25.107(a)(2) was revised to remove the reference to “takeoff decision speed” from the definition of V1. V1 is the speed by which the pilot has already made the decision to either continue or reject the takeoff and, if the latter, has initiated the first action to stop the airplane. 4.3.1.2.4 Section 25.109 was revised to add a requirement to determine accelerate-stop distances for wet runways. Additionally, the requirement for the AFM expansion to include two seconds of continued acceleration beyond V1, with the operating engines at takeoff power or thrust, as introduced by amendment 25-42, was replaced with a distance increment equivalent to two seconds at V1. Also, the text of § 25.109(a) was modified to clarify that the accelerate-stop distances must take into account the highest speed reached during the rejected takeoff maneuver, including, as applicable, speeds higher than V1. 4.3.1.2.5 Section 25.109(f) was added to permit credit for the use of reverse thrust in determining wet runway accelerate-stop distances (subject to the requirements of § 25.109(e)) and to explicitly deny reverse thrust credit for determining dry runway accelerate-stop distances. 4.3.1.2.6 Section 25.109(i) was added to require a maximum brake energy accelerate-stop test to be conducted with not more than 10 percent of the allowable brake wear range remaining on each individual wheel-brake assembly. 4-21 ### Page 57 **AC 25-7D CHG 1** 4.3.1.2.7 Section 25.735(h) was added to require the maximum rejected takeoff brake energy absorption capacity rating used during qualification testing to the applicable TSO to be based on the fully worn limit of the brake’s allowable wear range. Note: Section 25.735 was revised by amendment 25-107, which moved the rejected takeoff kinetic energy rating requirements into § 25.735(f). 4.3.1.2.8 Section 25.1533(a)(3) was revised to add runway surface condition (dry or wet) as a variable that must be accounted for in establishing minimum takeoff distances. Section 25.1533(a)(3) was also revised to allow wet runway takeoff distances on grooved and porous friction course (PFC) runways to be established as additional operating limitations, but approval to use these distances is limited to runways that have been designed, constructed, and maintained in a manner acceptable to the FAA Administrator. 4.3.2 Applicable Regulations. The applicable part 25 regulations are § 25.109, and the following:  Section 25.101(f) regarding airplane configuration and procedures.  Section 25.101(h) regarding pilot action time delay allowances.  Section 25.101(i) regarding worn brake stopping performance.  Section 25.105 regarding takeoff configuration and environmental and runway conditions.  Section 25.107(a)(l) and (a)(2) regarding V1 and VEF speed definitions.  Section 25.735 regarding brakes and braking systems.  Section 25.1301 regarding function and installation.  Section 25.1309 regarding equipment, systems, and installation.  Section 25.1533 regarding additional operating limitations—maximum takeoff weights and minimum takeoff distances.  Section 25.1583(h) regarding AFM operating limitations.  Section 25.1587 regarding AFM performance information. 4.3.3 Procedures: General. The following paragraphs provide guidance for accomplishing accelerate-stop flight tests and expanding the resulting data for the determination of AFM performance information. 4-22 ### Page 58 **AC 25-7D CHG 1** 4.3.4 Procedures: Accelerate-Stop Testing. The following guidance applies to turbine-powered airplanes with and without propellers. Guidance regarding flight testing applies only to dry runway accelerate-stop distances. Guidance for expanding the flight test data to determine AFM distances applies to both dry and wet runways, unless otherwise noted. Further guidance for determining wet runway accelerate-stop distances is provided in paragraph 4.3.7 of this AC. 4.3.4.1 To establish a distance that would be representative of the distance needed in the event of a rejected takeoff, where the first action to stop the airplane is taken at V1, a sufficient number of test runs should be conducted for each airplane configuration specified by the applicant. (For intermediate configurations, see paragraph 3.1 of this AC.) 4.3.4.2 The guidance outlined in paragraph 4.3.6 of this AC describes how to include allowances for any time delays, as required by § 25.101(h)(3), for the flightcrew to accomplish the rejected takeoff operating procedures. 4.3.4.3 Section 25.101(i) states that the accelerate-stop distances must be determined with all the airplane wheel-brake assemblies at the fully worn limit of their allowable wear range. The fully worn limit is defined as the maximum amount of wear allowed before the brake is to be removed from the airplane for overhaul. The allowable wear should be defined in terms of a linear dimension in the axial direction, which is typically determined by measuring the wear pin extension. 4.3.4.3.1 The only accelerate-stop test that must be conducted at a specific brake wear state is the maximum brake kinetic energy demonstration, which must use brakes that have no more than 10 percent of the allowable brake wear range remaining, as required by § 25.109(i). (See paragraph 4.3.5.3 of this AC). The remainder of the accelerate-stop tests may be conducted with the brakes in any wear state as long as a suitable combination of airplane and dynamometer tests is used to determine the accelerate-stop distances corresponding to fully worn brakes. For example, dynamometer testing may be used to determine whether there is a reduction in brake performance from the wear state used in the airplane tests to a fully worn brake. The airplane test data could then be adjusted analytically for this difference without additional airplane testing. 4.3.4.3.2 Either airplane-worn or mechanically-worn brakes (i.e., machined or dynamometer worn) may be used. If mechanically-worn brakes are used, it should be shown that they can be expected to provide similar results to airplane-worn brakes. This comparison can be based on service experience on the test brake or an appropriate equivalent brake, or on dynamometer wear test data when service data are unavailable. 4-23 ### Page 59 **AC 25-7D CHG 1** 4.3.4.4 Section 25.109(f)(1) denies credit for the use of reverse thrust as a decelerating means in determining the accelerate-stop distance for a dry runway. This provision applies to both turbine engine and propeller engine reverse thrust (but not to any drag resulting from the ground idle power setting for a propeller engine). Credit for the additional deceleration available from reverse thrust is permitted for wet runway accelerate-stop distances, provided the thrust reverser system is shown to be safe, reliable, capable of giving repeatable results, and does not require exceptional skill to control the airplane. (See paragraph 4.3.7.10 of this AC for guidance related to obtaining accelerate-stop performance credit for reverse thrust on wet runways.) 4.3.4.5 The accelerate-stop test runs should be conducted at weight/speed combinations that will provide an even distribution of test conditions over the range of weights, speeds, and brake energies for which takeoff data will be provided in the AFM. The effects of different airport elevations can be simulated at one airport elevation, provided the braking speeds employed are relevant for the range of airplane energies to be absorbed by the brakes. The limiting brake energy value in the AFM should not exceed the maximum demonstrated in these tests or the maximum for which the brake has been approved. (See paragraph 4.3.5 of this AC for further guidance related to tests and analyses for the demonstration of the maximum brake energy absorption capability.) 4.3.4.6 The V1 speeds used in the accelerate-stop tests need not correspond precisely to the AFM values for the test conditions since it may be necessary to increase or decrease the AFM V1 speed to investigate fully the energy range and weight envelope. 4.3.4.7 A total of at least six accelerate-stop flight tests should be conducted. Unless sufficient data are available for the specific airplane type showing how braking performance varies with weight, kinetic energy, lift, drag, ground speed, torque limit, etc., at least two tests should be conducted for each configuration when the same braking coefficient of friction is being claimed for multiple aerodynamic configurations. These tests should be conducted on smooth, hard surfaced, dry runways. 4.3.4.8 For approval of dispatch capability with anti-skid inoperative, nose wheel brakes or specified main wheel-brake(s) inoperative, automatic braking systems, etc., a full set of tests, as described in paragraph 4.3.4.7 above, should normally be conducted. A lesser number of tests may be accepted for “equal or better” demonstrations, to establish small increments or if adequate conservatism is used during testing. 4-24 ### Page 60 **AC 25-7D CHG 1** 4.3.4.9 Either ground or airborne instrumentation should include means to determine the horizontal distance time-history. 4.3.4.10 The wind speed and direction relative to the test runway should be determined and corrected to a height corresponding to the approximate height of the mean aerodynamic chord. (See paragraph 3.1 of this AC.) 4.3.4.11 The accelerate-stop tests should be conducted in the following configurations: 4.3.4.11.1 Heavy to light weight as required. 4.3.4.11.2 Most critical CG position. 4.3.4.11.3 Wing flaps in the takeoff position(s). 4.3.4.11.4 Tire pressure: before taxi and with cold tires, set to the highest value appropriate to the takeoff weight for which approval is being sought. 4.3.4.11.5 Engine idle power or thrust: set at the recommended upper limit for use on the ground or the effect of maximum ground idle power or thrust may be accounted for in data analyses. For maximum brake energy and fuse plug no-melt tests, data analysis may not be used in place of maximum ground idle power or thrust. 4.3.4.12 Engine power or thrust should be appropriate to each segment of the rejected takeoff and should include accounting for power or thrust decay rates (i.e., spin down) for failed or throttled back engines. 4.3.4.12.1 Turbojet Powered Airplanes. For AFM calculation purposes, the critical engine failure accelerate-stop data may be based on the failed engine spinning down to a windmilling condition. Note: If, due to the certification basis of the airplane, all-engine-accelerate-stop distances are not being considered, the one-engine-inoperative AFM distances should be based on the critical engine failing to maximum ground idle power or thrust rather than the windmilling condition. 4-25 ### Page 61 **AC 25-7D CHG 1** 4.3.4.12.2 The power or thrust from the operative engine(s) should be consistent with a throttle chop to maximum ground idle power or thrust. For determining the all-engines-operating dry runway accelerate-stop AFM distances, the stopping portion should be based on all engines producing maximum ground idle power or thrust (after engine spin down), as noted in paragraph 4.3.4.11.5 above. The accelerate-stop tests may be conducted with either concurrent or sequential throttle chops to idle power or thrust as long as the data are adjusted to take into account pilot reaction time, and any control, system, or braking differences (e.g., electrical or hydraulic/mechanical transients associated with an engine failing to a windmilling condition resulting in reduced braking effectiveness). Test data should also be analytically corrected for any differences between maximum ground idle power or thrust and the idle power or thrust level achieved during the test. For the criteria relating to reverse thrust credit for wet runway accelerate-stop distances, see paragraph 4.3.7.10 of this AC. 4.3.4.12.3 Turbopropeller-Powered Airplanes. For the one-engine-inoperative accelerate-stop distances, the critical engine’s propeller should be in the position it would normally assume when an engine fails and the power levers are closed. For dry runway one-engine-inoperative accelerate-stop distances, the high drag groundidle position of the operating engines’ propellers (defined by a pitch setting that results in not less than zero total thrust, i.e., propeller plus jet thrust, at zero airspeed) may be used provided adequate directional control is available on a wet runway and the related operational procedures comply with § 25.101(f) and (h). Wet runway controllability may either be demonstrated by using the guidance available in paragraph 4.3.7.10.6 of this AC at the appropriate power level, or adequate control can be assumed to be available at ground idle power if reverse thrust credit is approved for determining the wet runway accelerate-stop distances. For the all-engines-operating accelerate-stop distances on a dry runway, the high drag ground-idle propeller position may be used for all engines (subject to § 25.101(f) and (h)). For the criteria relating to reverse thrust credit for wet runway accelerate-stop distances, see paragraph 4.3.7.10 of this AC. 4.3.4.13 System transient effects (e.g., engine spin-down, brake pressure ramp-up, etc.) should be determined and properly accounted for in the calculation of AFM accelerate-stop distances. (See paragraph 4.3.6.10 of this AC.) 4-26 ### Page 62 **AC 25-7D CHG 1** 4.3.5 Procedures: Maximum Brake Energy Testing. The following paragraphs describe regulatory requirements and acceptable test methods for conducting an accelerate-stop test run to demonstrate the maximum energy absorption capability of the wheel brakes. 4.3.5.1 The maximum brake energy accelerate-stop demonstration should be conducted at not less than the maximum takeoff weight. It should be preceded by at least a 3-mile taxi with all engines operating at maximum ground idle power or thrust, including three full stops using normal braking. Following the maximum brake energy stop, it will not be necessary to demonstrate the airplane’s ability to taxi. 4.3.5.2 Section 25.735(f)(2) requires the maximum kinetic energy accelerate-stop absorption capability of each wheel, tire, and brake assembly to be determined. It also requires dynamometer testing to show that the wheel, brake, and tire assembly is capable of absorbing not less than this level of kinetic energy throughout the defined wear range of the brake. The calculation of maximum brake energy limited takeoff weights and speeds, for presentation in the AFM performance section, should, therefore, be based on the most critical wear range of the brake. 4.3.5.3 Section 25.109(i) requires a flight test demonstration of the maximum brake kinetic energy accelerate-stop distance to be conducted with not more than 10 percent of the allowable brake wear range remaining on each of the airplane wheel-brakes. The 10 percent allowance on the brake wear state is intended to ease test logistics and increase test safety, not to allow the accelerate-stop distance to be determined with less than fully worn brakes. If the brakes are not in the fully worn state at the beginning of the test, the accelerate-stop distance should be corrected as necessary to represent the stopping capability of fully worn brakes. 4.3.5.4 The maximum airplane brake energy allowed for dispatch should not exceed the value for which a satisfactory after-stop condition exists, or the value documented under the applicable TSO (or an acceptable equivalent), whichever value is less. A satisfactory after-stop condition is defined as one in which fires are confined to tires, wheels, and brakes, such that progressive engulfment of the rest of the airplane would not occur during the time of passenger and crew evacuation. The application of firefighting means or artificial coolants should not be required for a period of 5 minutes following the stop. 4.3.5.5 Landings are not an acceptable means for demonstrating the maximum rejected takeoff brake energy. Though permitted in the past, service experience has shown that methods used to predict brake and tire temperature increases that would have occurred during taxi and 4-27 ### Page 63 **AC 25-7D CHG 1** acceleration, as specified in paragraph 4.3.5.1 of this AC, were not able to accurately account for the associated energy increments. 4.3.6 Procedures: Accelerate-Stop Time Delays. 4.3.6.1 Section 25.101(h) requires allowance for time delays in the execution of procedures. Amendment 25-42 (effective March 1, 1978) amended the airworthiness standards to clarify and standardize the method of applying these time delays to the accelerate-stop transition period. Amendment 25-42 also added the critical engine failure speed, VEF, and clarified the meaning of V1 with relation to VEF. The preamble to amendment 25-42 states that “V1 is determined by adding to VEF (the speed at which the critical engine is assumed to fail) the speed gained with the critical engine inoperative during the time interval between the instant at which the critical engine is failed and the instant at which the test pilot recognizes and reacts to the engine failure, as indicated by the pilot’s application of the first retarding means during accelerate-stop tests.” Thus, it can be seen that V1 is not only intended to be at the end of the decision process, but it also includes the time it takes for the pilot to perform the first action to stop the airplane. (See appendix C of this AC for further discussion on the historical development of accelerate-stop time delays.) The purpose of the time delays is to allow sufficient time (and distance) for a pilot, in actual operations, to accomplish the procedures for stopping the airplane. The time delays are not intended to allow extra time for making a decision to stop as the airplane passes through V1. Since the typical transport category airplane requires three pilot actions (i.e., brakes-throttles-spoilers) to achieve the final braking configuration, amendment 25-42 defined a two-second time period, in § 25.109, to account for delays in activating the second and third deceleration devices. Amendment 25-92 (effective March 20, 1998) redefined and reinterpreted the application of that two-second delay time as a distance increment equivalent to two seconds at V1. No credit may be taken for system transient effects (e.g., engine spin-down, brake pressure ramp-up, etc.) in determining this distance. The following paragraphs provide guidance related to the interpretation and application of delay times to show compliance with the accelerate-stop requirements of amendment 25-92. 4.3.6.2 Figure 4-1 below presents a pictorial representation of the accelerate-stop time delays considered acceptable for compliance with § 25.101(h) as discussed above. 4-28 ### Page 64 AC 25-7D CHG 1 Figure 4-1. Accelerate-Stop Time Delays Event Demonstrated Time Delays VEF Flight Manual Expansion Time Delays Initiation o f Initiation of Initiation of Initiation of first pilot second pilot thirdpilot subsequent tac t 1 tact 2 tact 3 V1 tact 1 t act 2 tact 3 tact +1 sec. Distance equivalent to 2 sec. at V1 Engine failure action action action pilot actions tact 4 n 4.3.6.3 VEF is the calibrated airspeed selected by the applicant at which the critical engine is assumed to fail. The relationship between VEF and V1 is defined in § 25.107. 4.3.6.4 Δtact 1 is the demonstrated time interval between engine failure and initiation of the first pilot action to stop the airplane. This time interval is defined as beginning at the instant the critical engine is failed and ending when the pilot recognizes and reacts to the engine failure, as indicated by the pilot’s first action taken to stop the airplane during accelerate-stop tests. A sufficient number of demonstrations should be conducted using both applicant and FAA test pilots to assure that the time increment is representative and repeatable. The pilot’s feet should be on the rudder pedals, not the brakes, during the tests. For AFM data expansion purposes, in order to provide a recognition time increment that can be executed consistently in service, this time increment should be equal to the demonstrated time or one second, whichever is greater. If the airplane incorporates an engine failure warning light, the recognition time includes the time increment necessary for the engine to spool down to the point of warning light activation, plus the time increment from light “on” to the first pilot action to stop the airplane. 4.3.6.5 Δtact 2 is the demonstrated time interval between initiation of the first and second pilot actions to stop the airplane. 4-29 ### Page 65 **AC 25-7D CHG 1** 4.3.6.6 Δtact 3 is the demonstrated time interval between initiation of the second and third pilot actions to stop the airplane. 4.3.6.7 Δtact 4→n is the demonstrated time interval between initiation of the third and fourth (and any subsequent) pilot actions to stop the airplane. For AFM expansion, a one-second reaction time delay to account for inservice variations should be added to the demonstrated time interval between the third and fourth (and any subsequent) pilot actions. If a command is required for another crewmember to initiate an action to stop the airplane, a two-second delay, in lieu of the one-second delay, should be applied for each action. For automatic deceleration devices that are approved for performance credit for AFM data expansion, established systems actuation times determined during certification testing may be used without the application of the additional time delays required by this paragraph. 4.3.6.8 The sequence of pilot actions may be selected by the applicant, but it must match the sequence established for operation in service, as prescribed by § 25.101(f). If, on occasion, the specified sequence is not achieved during testing, the test need not be discarded; however, sufficient testing should be conducted to establish acceptable values of Δtact. 4.3.6.9 Section 25.109(a)(1)(iv) and (a)(2)(iii) require the one-engine-inoperative and all-engines-operating accelerate-stop distances, respectively, to include a distance increment equivalent to two seconds at V1. (Although the requirement for the distance increment equivalent to two seconds at V1 is explicitly stated in the “dry runway” criteria of § 25.109, it is also applied to the “wet runway” accelerate-stop distances by reference in § 25.109(b).) This distance increment is represented pictorially on the right side of the “Flight Manual Expansion Time Delays” presentation in figure 4-1 above, and in the speed versus distance plot of figure 4-2 below, on the following page. The two-second time period is only provided as a method to calculate the required distance increment and is not considered to be a part of the accelerate-stop braking transition sequence. Consequently, no credit for pilot actions, or engine and systems transient responses (e.g., engine spin-down) may be taken during this two-second time period. Similarly, the two-second time period may not be reduced for airplanes incorporating automated systems that decrease the number of pilot actions required to obtain the full braking configuration (e.g., automatic spoiler systems). 4-30 ### Page 66 AC 25-7D CHG 1 Figure 4-2. Accelerate-Stop Speed versus Distance DISTANCE 3rd Action 2nd Action V1 (1st Action) VEF AFM model must accurately represent demonstrated energy state of the airplane after V1. } Distance for 2 secs. at V1 SPEED 4.3.6.10 Section 25.109(a)(1)(ii) requires that any residual acceleration causing the airplane to exceed V1, while the airplane and its systems become stabilized in the braking configuration, be included in the accelerate-stop distance. The effects of system transients, such as engine spin-down, brake pressure ramp-up, spoiler actuation times, etc., should be accounted for in this time period. The area of interest is noted at the top of the graphical representation of the speed versus distance relationship in figure 4-2 above. 4.3.6.11 All-Engine Accelerate-Stop Distance. For the all-engines-operating accelerate-stop distance prescribed by § 25.109(a)(2), apply the demonstrated time intervals and associated delays, of paragraphs 4.3.6.5 through 4.3.6.7 of this AC, after the airplane has accelerated to V1. 4-31 ### Page 67 **AC 25-7D CHG 1** 4.3.6.12 Describe the procedures used to determine the accelerate-stop distance in the performance section of the AFM. 4.3.7 Procedures: Wet Runway Accelerate-Stop Distance. 4.3.7.1 The following guidance is provided for showing compliance with the requirements stated in § 25.109(b) through (d) for determining accelerate-stop distances applicable to wet runways. In general, the wet runway accelerate-stop distance is determined in a similar manner to the dry runway accelerate-stop distance. The only differences are in reflecting the reduced stopping force available from the wheel brakes on the wet surface and in provisions for performance credit for the use of reverse thrust as an additional decelerating means. The general method for determining the reduced stopping capability of the wheel brakes on a smooth wet runway is as follows: First, determine the maximum tire-toground wet runway braking coefficient of friction versus ground speed from the relationships provided in § 25.109(c)(1). Then, adjust this braking coefficient to take into account the efficiency of the anti-skid system. (See paragraph 4.3.7.3 of this AC for a definition of anti-skid efficiency. See paragraphs 4.3.7.5 and 4.3.7.6 for material on how to determine the wet runway anti-skid efficiency.) Next, determine the resulting braking force and adjust this force for the effect of the distribution of the normal load between braked and unbraked wheels at the most adverse CG position approved for takeoff, as prescribed by § 25.109(b)(2)(ii). (See paragraph 4.3.7.8 for a discussion of normal load distribution.) In accordance with § 25.109(b)(2)(i), apply further adjustments, if necessary, to ensure that the resulting stopping force attributed to the wheel brakes on a wet runway never exceeds (i.e., during the entire stop) the wheel brakes stopping force used to determine the dry runway accelerate-stop distance (under § 25.109(a)). Neither the dry runway brake torque limit nor the dry runway friction (i.e., anti-skid) limit should be exceeded. Alternative methods of determining the wet runway 4-32 ### Page 68 **AC 25-7D CHG 1** wheel brakes stopping force may be acceptable as long as that force does not exceed the force determined using the method just described. 4.3.7.2 Maximum Tire-to-Ground Braking Coefficient of Friction. The values specified in § 25.109(c)(1) were derived from data contained in Engineering Sciences Data Unit (ESDU) 71026, Frictional and Retarding Forces on Aircraft Types, Part II: Estimation of Braking Force, dated August 1981). The data in ESDU 71026 is a compilation from many different sources, including the National Aeronautics and Space Administration, the British Ministry of Aviation, and others. ESDU 71026 contains curves of wet runway braking coefficients versus speed for smooth and treaded tires at varying inflation pressures. These data are presented for runways of various surface roughness, including grooved and porous friction course runways. Included in the data presentation are bands about each of the curves, which represent variations in water depths from damp to flooded, runway surface texture within the defined texture levels, tire characteristics, and experimental methods. In defining the standard curves of wet runway braking coefficient versus speed that are prescribed by the equations in § 25.109(c)(1), the effects of the following variables were considered: tire pressure, tire tread depth, runway surface texture, and the depth of the water on the runway. 4.3.7.2.1 Tire Pressure. Lower tire pressures tend to improve the airplane’s stopping capability on a wet runway. The effect of tire pressure is taken into account by providing separate curves (equations) in § 25.109(c)(1) for several tire pressures. As stated in the rule, the tire pressure used to determine the maximum tire-to-ground braking coefficient of friction must be the maximum tire pressure approved for operation. Linear interpolation may be used for tire pressures other than those listed. 4.3.7.2.2 Tire Tread Depth. The degree to which water can be channeled out from under the tires significantly affects wet runway stopping capability. The standard curves of braking coefficient versus speed prescribed in § 25.109(c)(1) are based on a tire tread depth of 2 mm. This tread depth is consistent with tire removal and retread practices reported by airplane and tire manufacturers and tire retreaders. It is also consistent with FAA guidance provided in AC 121.195(d)-1A, regarding the tread depth for tires used in flight tests to determine operational landing distances on wet runways. Although operation with zero tread depth is not prohibited, it is unlikely that all of the tires on an airplane would be worn to the same extent. 4-33 ### Page 69 **AC 25-7D CHG 1** 4.3.7.2.3 Runway Surface Texture. ESDU 71026 groups runways into five categories. These categories are labeled “A” through “E,” with “A” being the smoothest and “C” the most heavily textured ungrooved runways. Categories “D” and “E” represent grooved and other open textured surfaces. Category A represents a very smooth texture (an average texture depth of less than 0.004 inches) and is not very prevalent in runways used by transport category airplanes. The majority of ungrooved runways fall into the category C grouping. The curves represented in § 25.109(c)(1) represent a texture midway between categories B and C. 4.3.7.2.4 Depth of Water on the Runway. Obviously, the greater the water depth, the greater the degradation in braking capability. The curves prescribed in § 25.109(c)(1) represent a well-soaked runway, but with no significant areas of standing water. 4.3.7.3 Anti-Skid System Efficiency. Section 25.109(c)(2) requires adjusting the maximum tire-to-ground braking coefficient determined in § 25.109(c)(1) to take into account the efficiency of the anti-skid system. The anti-skid system efficiency is defined as the relative capability of the anti-skid system to obtain the maximum friction available between the tire and the runway surface. It is expressed as either a percentage or a factor based on that percentage (e.g., 85 percent or 0.85). Applicants can either use one of the anti-skid efficiency values specified in § 25.109(c)(2) or derive the efficiency from flight tests on a wet runway. Regardless of which method is used, § 25.109(c)(2) requires that an appropriate level of flight testing be performed to verify that the anti-skid system operates in a manner consistent with the efficiency value used, and that the system has been properly tuned for operation on wet runways. 4.3.7.4 Classification of Types of Anti-Skid Systems. 4.3.7.4.1 The efficiency values specified in § 25.109(c)(2) are a function of the type of anti-skid system installed on the airplane. Three broad system types are identified in the rule: on/off, quasi-modulating, and fully modulating. These classifications represent evolving levels of technology and differing performance capabilities on dry and wet runways. The classification of anti-skid system types and the assigned efficiency values are based on information contained in Society of Automotive Engineers (SAE) Aerospace Information Report (AIR) 1739, Information on Anti-Skid Systems. 4-34 ### Page 70 **AC 25-7D CHG 1** 4.3.7.4.2 On/off systems are the simplest of the three types of anti-skid systems. For these systems, full-metered brake pressure (as commanded by the pilot) is applied until wheel locking is sensed. Brake pressure is then released to allow the wheel to spin back up. When the system senses that the wheel is accelerating back to synchronous speed (i.e., ground speed), full-metered pressure is again applied. The cycle of full pressure application/complete pressure release is repeated throughout the stop (or until the wheel ceases to skid with pressure applied). 4.3.7.4.3 Quasi-modulating systems attempt to continuously regulate brake pressure as a function of wheel speed. Typically, brake pressure is released when the wheel deceleration rate exceeds a pre-selected value. Brake pressure is re-applied at a lower level after a length of time appropriate to the depth of the skid. Brake pressure is then gradually increased until another incipient skid condition is sensed. In general, the corrective actions taken by these systems to exit the skid condition are based on a pre-programmed sequence rather than the wheel speed time history. 4.3.7.4.4 Fully modulating systems are a further refinement of the quasi-modulating systems. The major difference between these two types of anti-skid systems is in the implementation of the skid control logic. During a skid, corrective action is based on the sensed wheel speed signal, rather than a pre-programmed response. Specifically, the amount of pressure reduction or reapplication is based on the rate at which the wheel is going into or recovering from a skid. 4.3.7.4.5 In addition to examining the control system for the differences noted above, a time history of the response characteristics of the anti-skid system during a wet runway stop should be used to help identify the type of anti-skid system. Comparing the response characteristics between wet and dry runway stops can also be helpful. 4.3.7.4.6 Figure 4-3 and figure 4-4 below show an example of the response characteristics of a typical on-off system on both dry and wet runways, respectively. In general, the on-off system exhibits a cyclic behavior of brake pressure application until a skid is sensed, followed by the complete release of brake pressure to allow the wheel to spin back up. Full-metered pressure (as commanded by the pilot) is then re-applied, starting the cycle over again. The wheel speed trace exhibits deep and frequent skids (the troughs in the wheel speed trace), and the average wheel speed is significantly less than the synchronous speed (which is represented by the flat-topped portions of the wheel speed trace). Note that the skids are deeper and more frequent on a wet runway than on a dry runway. For the particular example shown in figure 4-3, the brake becomes torque-limited 4-35 ### Page 71 AC 25-7D CHG 1 toward the end of the dry runway stop, and it is unable to generate enough torque to cause further skidding. Figure 4-3. Anti-Skid System Response Characteristics: On-Off System on Dry Runway Figure 4-4. Anti-Skid System Response Characteristics: On-Off System on Wet Runway 4-36 ### Page 72 **AC 25-7D CHG 1** 4.3.7.4.7 The effectiveness of quasi-modulating systems can vary significantly depending on the slipperiness of the runway and the design of the particular control system. On dry runways, these systems typically perform very well; however, on wet runways their performance is highly dependent on the design and tuning of the particular system. An example of the response characteristics of one such system on dry and wet runways is shown in figure 4-5 and figure 4-6 below, respectively. On both dry and wet runways, brake pressure is released to the extent necessary to control skidding. As the wheel returns to the synchronous speed, brake pressure is quickly increased to a pre-determined level and then gradually ramped up to the full-metered brake pressure. On a dry runway, this type of response reduces the depth and frequency of skidding compared to an on-off system. However, on a wet runway, skidding occurs at a pressure below that at which the gradual ramping of brake pressure occurs. As a result, on wet runways the particular system shown in figure 4-6 operates very similarly to an on-off system. Figure 4-5. Anti-Skid System Response Characteristics: Quasi-Modulating System on Dry Runway 4-37 ### Page 73 AC 25-7D CHG 1 Figure 4-6 Anti-Skid System Response Characteristics: Quasi-Modulating System on Wet Runway 4.3.7.4.8 When properly tuned, fully modulating systems are characterized by much smaller variations in brake pressure around a fairly high average value. These systems can respond quickly to developing skids and are capable of modulating brake pressure to reduce the frequency and depth of skidding. As a result, the average wheel speed remains much closer to the synchronous wheel speed. Figure 4-7 below illustrates an example of the response characteristics of a fully modulating system on dry and wet runways. 4-38 ### Page 74 AC 25-7D CHG 1 Figure 4-7. Anti-Skid System Response Characteristics: Fully Modulating System 4-39 ### Page 75 **AC 25-7D CHG 1** 4.3.7.5 Demonstration of Anti-Skid System Operation when Using the Anti-Skid Efficiency Values Specified in § 25.109(c)(2). 4.3.7.5.1 If the applicant elects to use one of the anti-skid efficiency values specified in § 25.109(c)(2), a limited amount of flight testing must still be conducted, per § 25.109(c)(2), to demonstrate anti-skid system operation on a wet runway. This testing should be used to verify that the anti-skid system operates in a manner consistent with the type of anti-skid system declared by the applicant, and that the anti-skid system has been properly tuned for operation on wet runways. 4.3.7.5.2 A minimum of one complete stop, or equivalent segmented stops, should be conducted on a smooth (i.e., not grooved or porous friction course) wet runway at an appropriate speed and energy to cover the critical operating mode of the anti-skid system. Since the objective of the test is to observe the operation (i.e., cycling) of the anti-skid system, this test will normally be conducted at an energy level well below the maximum brake energy condition. 4.3.7.5.3 The section of the runway used for braking should be well soaked (i.e., not just damp), but not flooded. The runway test section should be wet enough to result in a number of cycles of anti-skid activity, but should not cause hydroplaning. 4.3.7.5.4 Before taxi and with cold tires, the tire pressure should be set to the highest value appropriate to the takeoff weight for which approval is being sought. 4.3.7.5.5 The tires and brakes should not be new but need not be in the fully worn condition. They should be in a condition considered representative of typical in-service operations. 4.3.7.5.6 Sufficient data should be obtained to determine whether the system operates in a manner consistent with the type of anti-skid system declared by the applicant, provide evidence that full brake pressure is being applied upstream of the anti-skid valve during the flight test demonstration, determine whether the anti-skid valve is performing as intended, and show that the anti-skid system has been properly tuned for a wet runway. Typically, the following parameters should be plotted versus time: 1. The speed of a representative number of wheels. 2. The hydraulic pressure at each brake (i.e., the hydraulic pressure downstream of the anti-skid valve or the electrical input to each anti-skid valve). 4-40 ### Page 76 **AC 25-7D CHG 1** 3. The hydraulic pressure at each brake metering valve (i.e., upstream of the anti-skid valve). 4.3.7.5.7 A qualitative assessment of anti-skid system response and airplane controllability should be made by the test pilot(s). In particular, pilot observations should confirm that: 1. Anti-skid releases are neither excessively deep nor prolonged; 2. The landing gear is free of unusual dynamics; and 3. The airplane tracks essentially straight, even though runway seams, water puddles, and wetter patches may not be uniformly distributed in location or extent. 4.3.7.6 Determination of a Specific Wet Runway Anti-Skid System Efficiency. 4.3.7.6.1 If the applicant elects to derive the anti-skid system efficiency from flight test demonstrations, sufficient flight testing, with adequate instrumentation, should be conducted to ensure confidence in the value obtained. An anti-skid efficiency of 92 percent (i.e., a factor of 0.92) is considered to be the maximum efficiency on a wet runway normally achievable with fully modulating digital anti-skid systems. 4.3.7.6.2 A minimum of three complete stops, or equivalent segmented stops, should be conducted on a wet runway at appropriate speeds and energies to cover the critical operating modes of the anti-skid system. Alternatively, if the operation and efficiency of the anti-skid system on a wet runway can be predicted by laboratory simulation data and validated by flight test demonstrations, a lesser number of stops may be acceptable. In this case, as many complete stops, or equivalent segmented stops, as necessary to present six independent anti-skid efficiency calculations should be conducted on a wet runway at appropriate speeds and energies to cover the critical operating modes of the anti-skid system. An independent anti-skid efficiency calculation can be presented for each stop for each independently controlled wheel or set of wheels. 4.3.7.6.3 Since the objective of the test is to determine the efficiency of the antiskid system, these tests will normally be conducted at energies well below the maximum brake energy condition. A sufficient range of speeds should be covered to investigate any variation of the anti-skid efficiency with speed. 4-41 ### Page 77 **AC 25-7D CHG 1** 4.3.7.6.4 The testing should be conducted on a smooth (i.e., not grooved or porous friction course) runway. If the applicant chooses to determine acceleratestop distances for grooved and porous friction course (PFC) surfaces under § 25.109(d)(2), testing should also be conducted on a grooved or porous friction course runway to determine the anti-skid efficiency value applicable to those surfaces. Other means for determining the anti-skid efficiency value for grooved and PFC surfaces may also be acceptable, such as using the efficiency value previously determined for smooth runways, if that value is shown to also be representative of or conservative for grooved and PFC runways. 4.3.7.6.5 The section of the runway used for braking should be well soaked (i.e., not just damp), but not flooded. The runway test section should be wet enough to result in a number of cycles of anti-skid activity but should not cause hydroplaning. 4.3.7.6.6 Before taxi and with cold tires, the tire pressure should be set to the highest value appropriate to the takeoff weight for which approval is being sought. 4.3.7.6.7 The tires and brakes should not be new but need not be in the fully worn condition. They should be in a condition considered representative of typical in-service operations. 4.3.7.6.8 A qualitative assessment of anti-skid system response and airplane controllability should be made by the test pilot(s). In particular, pilot observations should confirm that:  The landing gear is free of unusual dynamics; and  The airplane tracks essentially straight, even though runway seams, water puddles, and wetter patches may not be uniformly distributed in location or extent. 4.3.7.6.9 Two acceptable methods, referred to as the torque method and the wheel slip method, for determining the wet runway anti-skid efficiency value from wet runway stopping tests are described below. Other methods may also be acceptable if they can be shown to give equivalent results. The test instrumentation and data collection should be consistent with the method used. 4-42 ### Page 78 **AC 25-7D CHG 1** 4.3.7.6.10 Torque Method. 1. Under the torque method, the anti-ski d system efficiency is determined by comparing the energy absorbed by the brake during an actual wet runway stop to the energy that is determined by integrating, over the stopping distance, a curve defined by connecting the peaks of the instantaneous brake force curve. (See figure 4-8 below.) The energy absorbed by the brake during the actual wet runway stop is determined by integrating the curve of instantaneous brake force over the stopping distance. Figure 4-8. Instantaneous Brake Force and Peak Brake Force 2. Using data obtained from the wet runway stopping tests of paragraph 4.3.7.6 of this AC, instantaneous brake force can be calculated from the following relationship: Where: 3. For brake installations where measuring brake torque directly is impractical, torque may be determined from other parameters (e.g., brake pressure) if a suitable correlation is available. Wheel acceleration is obtained from the first derivative of wheel speed. Instrumentation recording rates and data analysis techniques for wheel speed and torque data should be well matched to the anti-skid response 4-43 ### Page 79 AC 25-7D CHG 1 characteristics to avoid introducing noise and other artifacts of the instrumentation system into the data. 4. Since the derivative of wheel speed is used in calculating brake force, smoothing of the wheel speed data is usually necessary to give good results. The smoothing algorithm should be carefully designed as it can affect the resulting efficiency calculation. Filtering or smoothing of the brake torque or brake force data should not normally be done. If conditioning is applied, it should be done in a conservative manner (i.e., result in a lower efficiency value) and should not misrepresent actual airplane/system dynamics. 5. Both the instantaneous brake force and the peak brake force should be integrated over the stopping distance. The anti-skid efficiency value for determining the wet runway accelerate-stop distance is the ratio of the instantaneous brake force integral to the peak brake force integral: Where: 6. The stopping distance is defined as the distance traveled during the specific wet runway stopping demonstration, beginning when the full braking configuration is obtained and ending at the lowest speed at which anti-skid cycling occurs (i.e., the brakes are not torque-limited), except that this speed need not be less than 10 knots. Any variation in the anti-skid efficiency with speed should also be investigated, which can be accomplished by determining the efficiency over segments of the total stopping distance. If significant variations are noted, this variation should be reflected in the braking force used to determine the accelerate-stop distances (either by using a variable efficiency or by using a conservative single value). 4.3.7.6.11 Wheel Slip Method. 1. At brake application, the tire begins to slip with respect to the runway surface (i.e., the wheel speed slows down with respect to the airplane’s ground speed). As the amount of tire slip increases, the brake force also increases until an optimal slip is reached. If the amount of slip continues to increase past the optimal slip, the braking force will decrease. 2. Using the wheel slip method, the anti- skid efficiency is determined by comparing the actual wheel slip measured during a wet runway stop to the optimal slip. Since the wheel slip varies significantly during the stop, sufficient wheel and ground speed data should be obtained to 4-44 ### Page 80 AC 25-7D CHG 1 determine the variation of both the actual wheel slip and the optimal wheel slip over the length of the stop. A sampling rate of at least 16 samples per second for both wheel speed and ground speed has been found to yield acceptable fidelity. 3. For each wheel and ground speed data point, the instantaneous antiskid efficiency value should be determined from the relationship shown in figure 4-9 below. Figure 4-9. Anti-Skid Efficiency–Wheel Slip Relationship Wheel Slip Ratio0 free rolling 1 locked wheel 1.0 0 Optimal Slip 0.5 Anti-Skid Efficiency    WSR   WSR  3 for WSR < OPS Efficiency = 1.5   0.5   OPS   OPS  WSR = OPS Efficiency = 1.0  1  WSRWSR > OPS Efficiency = 0.51+   1  OPS   Wheel speed where WSR = the wheel slip ratio = 1      Ground speed  and OPS is the optimal slip ratio 4. To determine the overall anti-skid efficiency value for use in calculating the wet runway accelerate-stop distance, the instantaneous anti-skid efficiencies should be integrated with respect to distance and divided by the total stopping distance: Where: 4-45 ### Page 81 **AC 25-7D CHG 1** 5. The stopping distance is defined as the distance traveled during the specific wet runway stopping demonstration, beginning when the full braking configuration is obtained and ending at the lowest speed at which anti-skid cycling occurs (i.e., the brakes are not torque-limited), except that this speed need not be less than 10 knots. Any variation in the anti-skid efficiency with speed should also be investigated, which can be accomplished by determining the efficiency over segments of the total stopping distance. If significant variations are noted, this variation should be reflected in the braking force used to determine the accelerate-stop distances (either by using a variable efficiency or by using a conservative single value). 6. The applicant should provide substantiation of the optimal wheel slip value(s) used to determine the anti-skid efficiency value. An acceptable method for determining the optimal slip value(s) is to compare time history plots of the brake force and wheel slip data obtained during the wet runway stopping tests. For brake installations where measuring brake force directly is impractical, brake force may be determined from other parameters (e.g., brake pressure) if a suitable correlation is available. For those skids where wheel slip continues to increase after a reduction in the brake force, the optimal slip is the slip value corresponding to the brake force peak. See figure 4-10 below for an example and note how both the actual wheel slip and the optimal wheel slip can vary during the stop. Figure 4-10. Substantiation of the Optimal Slip Value Brake Force Slip Optimal Ratio Wheel Time 4.3.7.7 For dispatch with an inoperative anti-skid system (if approved), the wet runway accelerate-stop distances should be based on an efficiency no 4-46 ### Page 82 AC 25-7D CHG 1 higher than that allowed by § 25.109(c)(2) for an on-off type of anti-skid system. The safety of this type of operation should be demonstrated by flight tests conducted in accordance with paragraph 4.3.7.5 of this AC. 4.3.7.8 Distribution of the Normal Load between Braked and Unbraked Wheels. In addition to taking into account the efficiency of the anti-skid system, § 25.109(b)(2)(ii) also requires adjusting the braking force for the effect of the distribution of the normal load between braked and unbraked wheels at the most adverse CG position approved for takeoff. The stopping force due to braking is equal to the braking coefficient multiplied by the normal load (i.e., weight) on each braked wheel. The portion of the airplane’s weight being supported by the unbraked wheels (e.g., unbraked nose wheels) does not contribute to the stopping force generated by the brakes. In accordance with § 25.21(a), this effect must be taken into account for the most adverse CG position approved for takeoff, considering any redistribution in loads that occur due to the dynamics of the stop. The most adverse CG position is the position that results in the least load on the braked wheels. 4.3.7.9 Grooved and Porous Friction Course (PFC) Runways. 4.3.7.9.1 Properly designed, constructed, and maintained grooved and PFC runways can offer significant improvements in wet runway braking capability. A conservative level of performance credit is provided by § 25.109(d) to reflect this performance improvement and to provide an incentive for installing and maintaining such surfaces. 4.3.7.9.2 In accordance with §§ 25.105(c) and 25.109(d), applicants may optionally determine the accelerate-stop distance applicable to wet grooved and PFC runways. These data would be included in the AFM in addition to the smooth runway accelerate-stop distance data. The braking coefficient for determining the accelerate-stop distance on grooved and PFC runways is defined in § 25.109(d) as either 70 percent of the braking coefficient used to determine the dry runway accelerate-stop distances, or a curve based on ESDU 71026 data and derived in a manner consistent with that used for smooth runways. In either case, the brake torque limitations determined on a dry runway may not be exceeded. 4.3.7.9.3 Using a simple factor applied to the dry runway braking coefficient is acceptable for grooved and PFC runways because the braking coefficient’s variation with speed is much lower on these types of runways. On smooth wet runways, the braking coefficient varies significantly with speed, which makes it inappropriate to apply a simple factor to the dry runway braking coefficient. 4-47 ### Page 83 **AC 25-7D CHG 1** 4.3.7.9.4 For applicants who choose to determine the grooved/PFC wet runway accelerate-stop distances in a manner consistent with that used for smooth runways, § 25.109(d)(2) provides the maximum tire-to-ground braking coefficient applicable to grooved and PFC runways. This maximum tire-to-ground braking coefficient must be adjusted for the anti-skid system efficiency, either by using the value specified in § 25.109(c)(2) appropriate to the type of anti-skid system installed, or by using a specific efficiency established by the applicant. As anti-skid system performance depends on the characteristics of the runway surface, a system that has been tuned for optimum performance on a smooth surface may not achieve the same level of efficiency on a grooved or porous friction course runway, and vice versa. Consequently, if the applicant elects to establish a specific efficiency for use with grooved or PFC surfaces, anti-skid efficiency testing should be conducted on a wet runway with such a surface, in addition to testing on a smooth runway. Means other than flight testing may be acceptable, such as using the efficiency previously determined for smooth wet runways, if that efficiency is shown to be representative of, or conservative for, grooved and PFC runways. Per § 25.109(b)(2)(ii), the resulting braking force for grooved/PFC wet runways must be adjusted for the effect of the distribution of the normal load between braked and unbraked wheels. This adjustment will be similar to that used for determining the braking force for smooth wet runways, except that the braking dynamics should be appropriate to the braking force achieved on grooved and PFC wet runways. Due to the increased braking force on grooved and PFC wet runways, an increased download on the nose wheel and corresponding reduction in the download on the main gear is expected. 4.3.7.9.5 In accordance with §§ 25.1533(a)(3) and 25.1583(h), grooved and PFC wet runway accelerate-stop distances may be established as operating limitations and be presented in the AFM, but approval to use these distances is limited to runways that have been designed, constructed, and maintained in a manner acceptable to the FAA Administrator. Airplane operators who wish to use the grooved or PFC runway accelerate-stop distances will need to determine that the design, construction, and maintenance aspects are acceptable for each runway for which such credit is sought. AC 150/5320-12C provides guidance relative to acceptable design, construction, and maintenance practices for grooved and PFC runway surfaces. 4.3.7.10 Reverse Thrust Performance Credit. In accordance with § 25.109(f), reverse thrust may not be used to determine the accelerate-stop distances for a dry runway. For wet runway accelerate-stop distances, however, § 25.109(f) allows credit for the stopping force provided by reverse thrust if the requirements of § 4-48 ### Page 84 **AC 25-7D CHG 1** 25.109(e) are met. In addition, the procedures associated with the use of reverse thrust, which § 25.101(f) requires the applicant to provide, must meet the requirements of § 25.101(h). The following criteria provide acceptable means of demonstrating compliance with these requirements: 4.3.7.10.1 Procedures for using reverse thrust during a rejected takeoff should be developed and demonstrated. These procedures should include all of the pilot actions necessary to obtain the recommended level of reverse thrust, maintain directional control and safe engine operating characteristics, and return the reverser(s), as applicable, to either the idle or the stowed position. These procedures need not be the same as those recommended for use during a landing stop, but should not result in additional hazards (e.g., cause a flameout or any adverse engine operating characteristics), nor should they significantly increase flightcrew workload or training needs. 4.3.7.10.2 It should be demonstrated that using reverse thrust during a rejected takeoff complies with the engine operating characteristics requirements of § 25.939. The reverse thrust procedures may specify a speed at which the reverse thrust is to be reduced to idle in order to maintain safe engine operating characteristics. 4.3.7.10.3 The time sequence for the actions necessary to obtain the recommended level of reverse thrust should be demonstrated by flight test. The time sequence used to determine the accelerate-stop distances should reflect the most critical case relative to the time needed to deploy the thrust reversers. For example, on some airplanes the outboard thrust reversers are locked out if an outboard engine fails. This safety feature prevents the pilot from applying asymmetric reverse thrust on the outboard engines, but it may also delay the pilot’s selection of reverse thrust on the operable reversers. In addition, if the selection of reverse thrust is the fourth or subsequent pilot action to stop the airplane (e.g., after manual brake application, thrust/power reduction, and spoiler deployment), a one-second delay should be added to the demonstrated time to select reverse thrust. (See figure 4-1 of this AC.) 4.3.7.10.4 The response times of the affected airplane systems to pilot inputs should be taken into account. For example, delays in system operation, such as thrust reverser interlocks that prevent the pilot from applying reverse thrust until the reverser is deployed, should be taken into account. The effects of transient response characteristics, such as reverse thrust engine spin-up, should also be included. 4.3.7.10.5 To enable a pilot of average skill to consistently obtain the recommended level of reverse thrust under typical in-service conditions, a lever position that incorporates tactile feedback (e.g., a detent or stop) should be 4-49 ### Page 85 AC 25-7D CHG 1 provided. If tactile feedback is not provided, a conservative level of reverse thrust should be assumed. 4.3.7.10.6 The applicant should demonstrate that exceptional skill is not required to maintain directional control on a wet runway with a ten-knot crosswind from the most adverse direction. For demonstration purposes, a wet runway may be simulated by using a nose wheel free to caster on a dry runway. Symmetric braking should be used during the demonstration, and both all-engines-operating and critical-engine-inoperative reverse thrust should be considered. The brakes and thrust reversers may not be modulated to maintain directional control. The reverse thrust procedures may specify a speed at which the reverse thrust is reduced to idle in order to maintain directional controllability. 4.3.7.10.7 Compliance with the requirements of §§ 25.901(b)(2), 25.901(c), 25.1309(b), and 25.1309(c) will be accepted as providing compliance with the “safe and reliable” requirements of §§ 25.101(h)(2) and 25.109(e)(1). 4.3.7.10.8 The number of thrust reversers used to determine the wet runway accelerate-stop distance data provided in the AFM should reflect the number of engines assumed to be operating during the rejected takeoff, along with any applicable system design features. The all-enginesoperating accelerate-stop distances should be based on all thrust reversers operating. The one-engine-inoperative accelerate-stop distances should be based on failure of the critical engine. For example, if the outboard thrust reversers are locked out when an outboard engine fails, the one-engine-inoperative accelerate stop distances can only include reverse thrust from the inboard engine thrust reversers. 4.3.7.10.9 For the engine failure case, it should be assumed that the thrust reverser does not deploy (i.e., no reverse thrust or drag credit for deployed thrust reverser buckets on the failed engine). 4.3.7.10.10 For approval of dispatch with one or more inoperative thrust reverser(s), the associated performance information should be provided either in the AFM or the master minimum equipment list (MMEL). 4.3.7.10.11 The effective stopping force provided by reverse thrust in each, or at the option of the applicant, the most critical takeoff configuration, should be demonstrated by flight test. (One method of determining the reverse thrust stopping force would be to compare unbraked runs with and without the use of thrust reversers.) Regardless of the method used to demonstrate the effective stopping force provided by reverse thrust, flight test demonstrations should be conducted using all of the stopping means on which the AFM wet runway accelerate-stop distances are based in order to substantiate the accelerate-stop distances and ensure that no adverse 4-50 ### Page 86 **AC 25-7D CHG 1** combination effects are overlooked. These demonstrations may be conducted on a dry runway. 4.3.7.10.12 For turbopropeller powered airplanes, the criteria of paragraphs 4.3.7.10.1 through 4.3.7.10.11 above remain generally applicable. Additionally, the propeller of the inoperative engine should be in the position it would normally assume when an engine fails and the power lever is closed. Reverse thrust may be selected on the remaining engine(s). Unless this selection is achieved by a single action to retard the power lever(s) from the takeoff setting without encountering a stop or lockout, it should be regarded as an additional pilot action for the purposes of assessing delay times. If this action is the fourth or subsequent pilot action to stop the airplane, a one-second delay should be added to the demonstrated time to select reverse thrust. 4.4 Takeoff Path—§ 25.111. 4.4.1 Section 25.111(a). 4.4.1.1 Explanation. The takeoff path requirements of § 25.111, and the reductions to that path required by § 25.115, are established so that performance data can be provided in the AFM for use in assessing vertical clearance of obstacles under the takeoff flight path. 4.4.1.2 Procedures. The height references in § 25.111 should be interpreted as geometric heights. Section 25.111(a) requires the actual takeoff path (from which the net takeoff flight path is derived) to extend to the higher of where the airplane is 1,500 feet above the takeoff surface or to the altitude at which the transition to en route configuration is complete and the final takeoff speed (VFTO) is reached. 4.4.2 Section 25.111(a)(1)—Takeoff Path Power/Thrust Conditions. 4.4.2.1 Explanation. In accordance with § 25.111(d), the takeoff path must be established either from continuous demonstrated takeoffs or by synthesis from segments. If determined from segments, it must be shown by continuous demonstrated takeoff to be conservative, in accordance with § 25.111(d)(4). 4.4.2.2 Procedures. 4.4.2.2.1 To be assured that the predicted takeoff path is representative of actual performance, construct the path using the power or thrust required by § 4-51 ### Page 87 **AC 25-7D CHG 1** 25.101(c). This requires, in part, that the power or thrust be based on the particular ambient atmospheric conditions that are assumed to exist along the path. The standard lapse rate for ambient temperature as specified in the 1962 U.S. Standard Atmosphere part 1 should be used for power or thrust determination associated with each pressure altitude during the climb. 4.4.2.2.2 In accordance with § 25.111(c)(4), the power or thrust up to 400 feet above the takeoff surface must represent the power or thrust available along the path resulting from the power lever setting established during the initial ground roll in accordance with AFM procedures. This resulting power or thrust may be less than that available from the rated inflight setting schedule. 4.4.2.2.3 A sufficient number of takeoffs, to at least the altitude above the takeoff surface scheduled for V2 climb, should be made to establish the power or thrust lapse for a fixed power or thrust lever setting. An analysis may be used to account for various engine bleeds (e.g., ice protection, air conditioning, etc.) and for electrical, pneumatic, and mechanical power extraction. In some airplanes, the power or thrust growth characteristics are such that less than full rated power or thrust is used for AFM takeoff power/thrust limitations and performance. This is to preclude engine limitations from being exceeded during the takeoff climbs to 400 feet above the takeoff surface. 4.4.2.2.4 Engine power or thrust lapse with speed and altitude during the takeoff and climb, at fixed power or thrust lever settings), can be affected by takeoff pressure altitude. 4.4.2.2.5 Most turbine engines are sensitive to crosswind or tailwind conditions, when setting takeoff power under static conditions, and may stall or surge. To preclude this problem, it is acceptable to establish a rolling takeoff power or thrust setting procedure, provided the AFM takeoff field length and the takeoff power or thrust setting charts are based on this procedure. Demonstrations and analyses have been accepted in the past showing a negligible difference in distance between static and rolling takeoffs. A typical test procedure is as follows: 1. After stopping on the runway, set an intermediate power or thrust on all engines (power setting selected by applicant). 2. Release brakes and advance power or thrust levers. 3. Set target power or thrust setting as rapidly as possible prior to reaching 60 to 80 knots. 4. No adverse engine operating characteristics should exist after completion of the power setting through the climb to 1,500 feet 4-52 ### Page 88 **AC 25-7D CHG 1** above the airport and attainment of the en route configuration. Tests should be conducted to determine if any engine operating problems exist for takeoffs conducted throughout the altitude range for which takeoff operations are to be scheduled in the AFM. 4.4.2.2.6 If the applicant wishes to use a different procedure, it should be evaluated and, if found acceptable, the procedure should be reflected in the AFM. 4.4.3 Section 25.111(a)(2)—Engine Failure. 4.4.3.1 Explanation. 4.4.3.1.1 Since the regulations cannot dictate what type of engine failures may actually occur, it could be assumed that the engine failure required by the regulation occurs catastrophically. Such a failure would cause the power or thrust to drop immediately, with the associated performance going from all-engines-operating to one-engine-inoperative at the point of engine failure. 4.4.3.1.2 This conservative rationale notwithstanding, there is a basis for assuming that the failed engine power or thrust will not decay immediately. Unlike reciprocating engines, the locking-up of a jet engine fan without causing the engine to separate from the airplane is highly unlikely. Separation of the engine or fan, or fan disintegration, would remove weight and/or the ram drag included in the engine inoperative performance, providing compensation for the immediate thrust loss. 4.4.3.1.3 With these considerations, it may be acceptable to use the transient power or thrust as the failed engine spools down at VEF. The power or thrust time history used for data reduction and expansion should be substantiated by test results. 4.4.3.1.4 In the case of propeller-driven airplanes, consideration should also be given to the position of the failed engine’s propeller during the engine failure. These airplanes typically incorporate an automatic system to drive the propeller to a low drag position when an engine fails. The loss of power in this case will be much more sudden than the turbojet engine spooldown described above. 4.4.3.2 Procedures. 4.4.3.2.1 For turbojet powered airplanes, if transient thrust credit is used during engine failure in determining the one-engine-inoperative AFM takeoff performance, sufficient tests should be conducted using actual fuel cuts to establish the thrust decay as contrasted to idle engine cuts. For derivative programs not involving a new or modified engine type (i.e., a modification 4-53 ### Page 89 **AC 25-7D CHG 1** that would affect thrust decay characteristics), fuel cuts are unnecessary if thrust decay characteristics have been adequately substantiated. 4.4.3.2.2 For propeller driven airplanes, the use of fuel cuts can be more important in order to ensure that the takeoff speeds and distances are obtained with the critical engine’s propeller attaining the position it would during a sudden engine failure. The number of tests that should be conducted using fuel cuts, if any, depends on the correlation obtained with the idle cut data and substantiation that the data analysis methodology adequately models the effects of a sudden engine failure. 4.4.4 Section 25.111(a)(3)—Airplane Acceleration. 4.4.4.1 Explanation. None. 4.4.4.2 Procedures. None. 4.4.5 Section 25.111(b)—Airplane Rotation and Gear Retraction. 4.4.5.1 Explanation. The rotation speed, VR, is intended to be the speed at which the pilot initiates action to raise the nose wheel(s) off the ground during the acceleration to V2. Consequently, the takeoff path, determined in accordance with § 25.111(a) and (b), should assume that pilot action to raise the nose wheel(s) off the ground will not be initiated until the speed VR has been reached. 4.4.5.2 Procedures. The time between liftoff and initiation of gear retraction during takeoff distance demonstrations should not be less than that necessary to establish an indicated positive rate of climb plus one second. For the purposes of flight manual expansion, the average demonstrated time delay between liftoff and initiation of gear retraction may be assumed; however, this value should not be less than 3 seconds. 4.4.6 Section 25.111(c)(1)—Takeoff Path Slope. 4.4.6.1 Explanation. 4.4.6.1.1 Establishing a horizontal segment as part of the takeoff flight path is considered to be acceptable, per § 25.115(c), for showing compliance with the positive slope required by § 25.111(c)(1). 4-54 ### Page 90 **AC 25-7D CHG 1** 4.4.6.1.2 The net takeoff flight path is the flight path used to determine compliance with the airplane obstacle clearance requirements of the applicable operating rules. Section 25.115(b) states the required climb gradient reduction to be applied throughout the flight path for the determination of the net flight path, including the level flight acceleration segment. Rather than decreasing the level flight path by the amount required by § 25.115(b), § 25.115(c) allows the airplane to maintain a level net flight path during acceleration, but with a reduction in acceleration equal to the gradient decrement required by § 25.115(b). By this method, the altitude reduction is exchanged for increased distance to accelerate in the level flight portion of the net takeoff path. 4.4.6.2 Procedures. The level acceleration segment in the AFM net takeoff profile should begin at the horizontal distance along the takeoff flight path where the actual airplane height, without the gradient reductions of § 25.115(b), reaches the AFM specified acceleration height. 4.4.7 Section 25.111(c)(2)—Takeoff Path Speed. 4.4.7.1 Explanation. 4.4.7.1.1 It is intended that the airplane be flown at a constant indicated airspeed to at least 400 feet above the takeoff surface. This speed must meet the constraints on V2 of § 25.107(b) and (c). 4.4.7.1.2 The specific wording of § 25.111(c)(2) should not be construed to imply that above 400 feet the airspeed may be reduced below V2, but instead that acceleration may be commenced. 4.4.7.2 Procedures. 4.4.7.2.1 For those airplanes that take advantage of reduced stall speeds at low pressure altitude, the scheduling of V2 should not be factored against the stall speed obtained at the takeoff surface pressure altitude. Such a procedure would result in a reduced stall speed margin during the climb, which would be contrary to the intent of § 25.107(b). 4.4.7.2.2 For those airplanes mentioned in paragraph 4.4.7.1 above, the V2 should be constrained, in addition to the requirements of § 25.107(b) and (c), by the stall speed 1,500 feet above the takeoff surface. Weight reduction along the takeoff path, due to fuel burn, may be considered in the calculation of the stall speed ratios, provided it is well established. However, many applicants have measured stall speeds at 10,000 to 15,000 feet, which provides a conservative stall margin at lower takeoff field pressure altitudes. 4-55 ### Page 91 **AC 25-7D CHG 1** 4.4.8 Section 25.111(c)(3)—Required Gradient. 4.4.8.1 Explanation. None. 4.4.8.2 Procedures. None. 4.4.9 Section 25.111(c)(4)—Configuration Changes. 4.4.9.1 Explanation. 4.4.9.1.1 The intent of this requirement is to permit only those crew actions that are conducted routinely to be used in establishing the one-engine-inoperative takeoff path. The power/thrust levers may only be adjusted early during the takeoff roll and then left fixed until at least 400 feet above the takeoff surface. 4.4.9.1.2 Simulation studies and accident investigations have shown that when high workload occurs in the cockpit, as with an engine failure during takeoff, the crew might not advance the power/thrust on the operative engines, even if the crew knows the operative engines have been set at reduced power/thrust and a power/thrust increase is needed for terrain avoidance. This same finding applies to manually feathering a propeller. The landing gear may be retracted, however, as this is accomplished routinely once a positive rate of climb is observed. Also, automatic propeller feathering is specifically allowed by the rule. Guidance related to performance credit for automatic propeller feathering is provided in paragraph 4.4.9.2 below. 4.4.9.1.3 Although performance credit for pilot action to increase power/thrust below 400 feet above the takeoff surface is not permitted, performance credit is allowed for an automatic power advance. ATTCS are addressed in § 25.904, and the related performance requirements are described in appendix I to part 25. 4.4.9.2 Procedures. 4.4.9.2.1 Propeller pitch setting generally has a significant effect on minimum control speeds and airplane drag. The magnitude of these effects is such that continued safe flight may not be possible should a combined failure of an engine and its automatic propeller feathering system occur at a takeoff speed based on operation of that system. When this is the case, § 25.1309(b)(1) requires the probability of this combined failure to be extremely improbable (on the order of 10-9 per flight hour). 4-56 ### Page 92 **AC 25-7D CHG 1** 4.4.9.2.2 In determining the combined engine and automatic propeller drag reduction system failure rate of paragraph 4.4.9.2.1 above, the engine failure rate should be substantiated. Notwithstanding the in-service engine failure rate, the failure of the automatic propeller drag reduction system should be shown not to exceed 10-4 per flight hour. 4.4.9.2.3 The automatic propeller feathering system should be designed so that it will automatically be disabled on the operating engine(s) following its activation on the failed engine. The probability of an unwanted operation of the automatic propeller drag reduction system on an operating engine, following its operation on the failed engine, should be shown to be extremely improbable (on the order of 10-9 per flight hour). 4.4.9.2.4 If performance credit is given for operation of the automatic propeller feathering system in determining the takeoff distances, the propeller of the failed engine must also be assumed to be in the reduced drag position in determining the accelerate-stop distances. 4.4.9.2.5 The limitations section of the AFM should require the flightcrew to perform a functional check of the automatic propeller feathering system. The frequency with which flightcrew functional checks and ground maintenance checks must be performed should be considered in the evaluation of the system reliability. 4.4.9.2.6 Clear annunciations should be provided to the flightcrew to indicate the following when:  The automatic propeller drag reduction system is “ARMED”; and  A malfunction exists in the automatic propeller feathering system. 4.4.10 Section 25.111(d)—Takeoff Path Construction. 4.4.10.1 Explanation. This regulation should not be construed to mean that the takeoff path must be constructed entirely from a continuous demonstration or entirely from segments. To take advantage of ground effect, typical AFM takeoff paths use a continuous takeoff path from VLOF to the gear up point, covering the range of thrust-to-weight ratios. From that point free air performance, in accordance with § 25.111(d)(2), is added segmentally. This methodology may yield an AFM flight path that is steeper with the gear down than up. 4.4.10.2 Procedures. The AFM should include the procedures necessary to achieve this performance. 4-57 ### Page 93 **AC 25-7D CHG 1** 4.4.11 Section 25.111(d)(1)—Takeoff Path Segment Definition. 4.4.11.1 Explanation. None. 4.4.11.2 Procedures. None. 4.4.12 Section 25.111(d)(2)—Takeoff Path Segment Conditions. 4.4.12.1 Explanation. Section 25.111(d)(2) states: “The weight of the airplane, the configuration, and the power or thrust must be constant throughout each segment and must correspond to the most critical condition prevailing in the segment.” The intent is that, for simplified analysis, the performance is based on that value available at the most critical point in time during the segment, not that the individual variables (weight, approximate power or thrust setting, etc.) are each picked at their most critical value and then combined to produce the performance for the segment. 4.4.12.2 Procedures. The performance during the takeoff path segments should be obtained using one of the following methods: 4.4.12.2.1 The critical level of performance as explained in paragraph 4.4.12.1 above; 4.4.12.2.2 The average performance during the segment; or 4.4.12.2.3 The actual performance variation during the segment. 4.4.13 Section 25.111(d)(3)—Segmented Takeoff Path Ground Effect. 4.4.13.1 Explanation. This requirement does not intend the entire flight path to necessarily be based upon out-of-ground-effect performance simply because the continuous takeoff demonstrations have been broken into sections for data reduction expediency. For example, if the engine inoperative acceleration from VEF to VR is separated into a power or thrust decay portion and a windmilling drag portion, the climb from 35 feet to gear up does not necessarily need to be based upon out-of-ground-effect performance. (Also, see the explanation of § 25.111(d) in paragraph 4.4.10.1 of this AC.) 4-58 ### Page 94 **AC 25-7D CHG 1** 4.4.13.2 Procedures. None. 4.4.14 Section 25.111(d)(4)—Segmented Takeoff Path Check. 4.4.14.1 Explanation. None. 4.4.14.2 Procedures. If the construction of the takeoff path from brake release to out-of-groundeffect contains any portions that have been segmented (e.g., airplane acceleration segments with all-engines-operating and one-engine-inoperative), the path should be checked by continuous demonstrated takeoffs. A sufficient number of these, employing the AFM established takeoff procedures and speeds and covering the range of thrust-to-weight ratios, should be made to ensure the validity of the segmented takeoff path. The continuous takeoff data should be compared to takeoff data calculated by AFM data procedures but using test engine thrusts and test speeds. 4.4.15 Section 25.111(e)—Flight Path with Standby Power Rocket Engines. [Reserved] 4.5 Takeoff Distance and Takeoff Run—§ 25.113. 4.5.1 Takeoff Distance on a Dry Runway—§ 25.113(a). 4.5.1.1 The takeoff distance on a dry runway is the longer of the two distances described in paragraphs 4.5.1.1.1 and 4.5.1.1.2 below. The distances indicated below are measured horizontally from the main landing gears at initial brake release to that same point on the airplane when the lowest part of the departing airplane is 35 feet above the surface of the runway. 4.5.1.1.1 The distance measured to 35 feet with a critical engine failure occurring at VEF as shown in figure 4-11 below. 4-59 ### Page 95 AC 25-7D CHG 1 Figure 4-11. Takeoff Distance on a Dry Runway: Critical Engine Fails at VEF 4.5.1.1.2 One hundred fifteen percent of the distance measured to the 35 feet height above the takeoff surface with all-engines-operating as shown in figure 4-12 below. In establishing the all-engines-operating takeoff distance, § 25.113(a)(2) requires the distance to be “...determined by a procedure consistent with § 25.111” (Takeoff Path). The interpretation of this statement is that the all-engines-operating takeoff distance should:  Be based on the airplane reaching a speed of V2 before it is 35 feet above the takeoff surface; and  Be consistent with the achievement of a smooth transition to the steady initial climb speed at a height of 400 feet above the takeoff surface. Figure 4-12. Takeoff Distance: All-Engines-Operating 4.5.1.2 In accordance with § 25.101(f), the takeoff distance must be based on the procedures established for operation in service. 4-60 ### Page 96 **AC 25-7D CHG 1** 4.5.2 Takeoff Distance on a Wet Runway—§ 25.113(b). 4.5.2.1 The takeoff distance on a wet runway is the longer of the takeoff distance on a dry runway (using the dry runway V1 speed), determined in accordance with paragraphs 4.5.1.1.1 and 4.5.1.1.2 of this AC, or the distance on a wet runway using a reduced height at the end of the takeoff distance (and the wet runway V1 speed) as described in paragraph 4.5.2.2 below. 4.5.2.2 The takeoff distance on a wet runway is determined as the horizontal distance the main landing gear travels from brake release to the point where the lowest part of the airplane is 15 feet above the takeoff surface. The airplane must attain a height of 15 feet above the takeoff surface in a manner that will allow V2 to be achieved before reaching a height of 35 feet above the takeoff surface in accordance with § 25.113(b)(2) and as shown in figure 4-13 below. Figure 4-13. Takeoff Distance on a Wet Runway: Critical Engine Fails at VEF 4.5.3 Takeoff Run—§ 25.113(c). 4.5.3.1 The concept of takeoff run was introduced by SR-422A to allow credit for a portion of the airborne part of the takeoff distance to be flown over a clearway. (See paragraph 4.5.3.3 of this AC for a definition of clearway.) The takeoff run is the portion of the takeoff distance that must take place on or over the runway in accordance with the applicable operating rules. If there is no clearway, the takeoff run is equal to the takeoff distance. If there is a clearway, the takeoff run is the longer described in paragraph 4.5.3.1.1 or 4.5.3.1.2 below. These distances are measured as described in paragraph 4.5.1.1 for § 25.113(a). When using a clearway to determine the takeoff run, no more than one half of the air distance from VLOF to V35 may be flown over the clearway. 4-61 ### Page 97 **AC 25-7D CHG 1** 4.5.3.1.1 The takeoff run is the distance from the start of the takeoff roll to the midpoint between liftoff and the point at which the airplane attains a height of 35 feet above the takeoff surface, with a critical engine failure occurring at VEF, as shown in figure 4-14 below. For takeoff on a wet runway, the takeoff run is equal to the takeoff distance (i.e., there is no clearway credit allowed on a wet runway). Figure 4-14. Takeoff Run: Critical Engine Fails at VEF 4.5.3.1.2 One hundred fifteen percent of the distance from the start of the takeoff roll to the mid-point between liftoff and the point at which the airplane attains a height of 35 feet above the takeoff surface, with all engines operating, as shown in figure 4-15 of this AC. In establishing the all-engines-operating takeoff run, § 25.113(c)(2) requires the distance to be “...determined by a procedure consistent with § 25.111” (Takeoff Path). The interpretation of that statement is that the all-engines-operating takeoff run should:  Be based on the airplane reaching a speed of V2 before it is 35 feet above the takeoff surface; and  Be consistent with the achievement of a smooth transition to the steady initial climb speed at a height of 400 feet above the takeoff surface. 4-62 ### Page 98 AC 25-7D CHG 1 Figure 4-15. Takeoff Run: All-Engines-Operating 4.5.3.2 There may be situations where the takeoff run may be longer for the one-engine-inoperative condition (paragraph 4.5.3.1.1 of this AC) while the takeoff distance is longer for the all-engines-operating condition (paragraph 4.5.3.1.2 of this AC), or vice versa. Therefore, both conditions should always be considered. 4.5.3.3 Clearway is defined in 14 CFR part 1 as a plane extending from the end of the runway with an upward slope not exceeding 1.25 percent, above which no object nor any terrain protrudes. For the purpose of establishing takeoff distances and the length of takeoff runs, the clearway is considered to be part of the takeoff surface extending with the same slope as the runway, and the 35 feet height should be measured from that surface. (See figure 4-16 of this AC.) Figure 4-16. Clearway Profiles 4-63 ### Page 99 **AC 25-7D CHG 1** 4.6 Takeoff Flight Path—§ 25.115. 4.6.1 Takeoff Flight Path—§ 25.115(a). 4.6.1.1 Explanation. The takeoff flight path begins at the end of the takeoff distance and at a height of 35 feet above the takeoff surface. The takeoff flight path ends when the airplane’s actual height is the higher of 1,500 feet above the takeoff surface or at an altitude at which the en route configuration and final takeoff speed have been achieved. (See paragraph 4.4 of this AC for additional discussion.) Section 25.115(a) states that the takeoff “shall be considered to begin 35 feet above the takeoff surface,” recognizing that in the case of a wet runway the airplane’s actual height will only be 15 feet. This wording allows the same takeoff flight path determined under § 25.115 for a dry runway takeoff to also be used for a wet runway takeoff. For takeoffs from wet runways, the actual airplane height will be 20 feet lower than the takeoff flight path determined under § 25.115 for takeoff from a dry runway. Therefore, the airplane will be 20 feet closer vertically to obstacles after taking off from a wet runway compared to taking off from a dry runway. 4.6.1.2 Procedures. See figure 4-17 of this AC. 4.6.1.2.1 In figure 4-17, the final takeoff segment will usually begin with the airplane in the en route configuration and with maximum continuous power or thrust, but it is not required that these conditions exist until the end of the takeoff path when compliance with § 25.121(c) is shown. The time limit on takeoff power or thrust cannot be exceeded. 4.6.1.2.2 In figure 4-17, path 1 depicts a flight path based on a minimum 400-foot level-off for acceleration and flap retraction following the second segment climb portion of the flight path. Path 2 depicts the upper limit of the takeoff flight path following an extended second segment. Depending on obstacle clearance needs, the second segment may be extended to a height of more than 1500 feet above the takeoff surface. 4-64 ### Page 100 AC 25-7D CHG 1 Figure 4-17. Takeoff Segments and Nomenclature 4.6.2 Net Takeoff Flight Path—§ 25.115(b) and (c). 4.6.2.1 Explanation. 4.6.2.1.1 The net takeoff flight path is the actual flight path diminished by a gradient of 0.8 percent for two-engine airplanes, 0.9 percent for three-engine airplanes, and 1.0 percent for four-engine airplanes. 4.6.2.1.2 For the level flight acceleration segment, these prescribed gradient reductions may be applied as an equivalent reduction in acceleration in lieu of reduction in net flight path. (See paragraph 4.4.6 of this AC for additional discussion.) 4.6.2.2 Procedures. See figure 4-18 below. 4-65 ### Page 101 AC 25-7D CHG 1 Figure 4-18. Net Takeoff Flight Path 4.7 Climb: General—§ 25.117. 4.7.1 Explanation. This section states the climb requirements of §§ 25.119 and 25.121 must be complied with at each weight, altitude, and ambient temperature within the operational limits established for the airplane and with the most unfavorable CG for each configuration. The effects of changes in the airplane’s true airspeed during a climb at the recommended constant indicated (or calibrated) climb speed should be taken into account when showing compliance with these climb requirements. 4.7.2 Procedures. None. 4.8 Landing Climb: All-Engines-Operating—§ 25.119. 4.8.1 Explanation. Section 25.119(a) states that the engines are to be set at the power or thrust that is available 8 seconds after starting to move the power or thrust controls from the minimum flight idle position to the go-around power or thrust setting. Use the procedures given below for the determination of this maximum power or thrust for showing compliance with the climb requirements of § 25.119. 4.8.2 Procedures. 4.8.2.1 The engines should be trimmed to the low side of the idle trim band, if applicable, as defined in the airplane maintenance manual. The effect of any variation in the idle fuel flow schedule for engines with electronic fuel 4-66 ### Page 102 AC 25-7D CHG 1 controllers is typically negligible (but any such claim should be adequately substantiated). 4.8.2.2 At the most adverse test altitude, not to exceed the maximum field elevation for which certification is sought plus 1,500 feet, and with the most adverse bleed configuration expected in normal operations, stabilize the airplane in level flight with symmetrical power or thrust on all engines, landing gear down, flaps in the landing position, at a speed of VREF. Retard the throttle(s) of the test engine(s) to flight idle and determine the time needed to reach a stabilized RPM, as defined below, for the test engine(s) while maintaining level flight or the minimum rate of descent obtainable with the power or thrust of the remaining engine(s) not greater than maximum continuous thrust (MCT). Engine flight idle RPM is considered to be stabilized when the initial rapid deceleration of all rotors is completed. This has usually been 8-20 seconds. This can be determined in the cockpit as the point where rapid movement of the tachometer ceases. For some airplanes it may be desirable to determine the deceleration time from plots of RPM versus time. 4.8.2.3 For the critical air bleed and power extraction configuration, stabilize the airplane in level flight with symmetric power or thrust on all engines, landing gear down, flaps in the landing position, at a speed of VREF, simulating the estimated minimum climb-limited landing weights at an altitude sufficiently above the selected test altitude so that time to descend to the test altitude with the throttles closed equals the appropriate engine RPM stabilization time determined in paragraph 4.8.2.2 above. Retard the throttles to the flight idle position and descend at VREF to approximately the test altitude. When the appropriate time has elapsed, rapidly advance the power or thrust controls to the go-around power or thrust setting. The power or thrust controls may first be advanced to the forward stop and then retarded to the go-around power or thrust setting. At the applicant’s option, additional less critical bleed configurations may be tested. 4.8.2.4 The power or thrust that is available 8 seconds after starting to move the power or thrust controls from the minimum flight idle position, in accordance with paragraph 4.8.2.3 above, is the maximum permitted for showing compliance with the landing climb requirements of § 25.119(a), and Section 4T.119(a) of SR-422B (see appendix D of this AC) for each of the bleed and power extraction combinations tested in accordance with paragraph 4.8.2.3 above. Unless AFM performance data are presented for each specific bleed and power extraction level, the AFM performance data should be based on the power or thrust obtained with the most critical power extraction level. 4.8.2.5 For airplanes equipped with autothrottles that will be used for approach and go-around, the effect of using the autothrottle to set go-around power 4-67 ### Page 103 **AC 25-7D CHG 1** or thrust should be determined. One way to do this would be to complete the test procedure given in paragraph 4.8.2.3, except that the airplane should be stabilized on -3° approach path, nominal power extraction, at the weight that gives the lowest power or thrust for that condition. The autothrottle should then be used to increase the power or thrust to the go-around power or thrust setting. The power or thrust used to show compliance with § 25.119(a) should be the lesser of: 4.8.2.5.1 The power or thrust that is available 8 seconds after selection of go-around power or thrust using the autothrottle; or 4.8.2.5.2 The power or thrust determined under paragraph 4.8.2.4 of this AC. 4.9 Climb: One-Engine-Inoperative—§ 25.121. 4.9.1 Explanation. Section 25.121 contains one-engine-inoperative climb gradient capability requirements for the first, second, and final takeoff segments as well as for approach. 4.9.2 Procedures. 4.9.2.1 Two methods for establishing one-engine-inoperative climb performance follow: 4.9.2.1.1 Reciprocal heading climbs are conducted at several thrust-to-weight conditions from which the performance for the AFM is extracted. These climbs are flown with the wings nominally level. Control forces should be trimmed out as much as possible, except for the takeoff climbs with gear extended where the takeoff trim settings are to be retained to represent an operationally realistic drag level. Reciprocal climbs may not be necessary if inertial corrections (or another equivalent means) are applied to account for wind gradients. 4.9.2.1.2 Drag polars and one-engine-inoperative yaw drag data are obtained for expansion into AFM climb performance. These data are obtained with the wings nominally level. Reciprocal heading check climbs are conducted to verify the predicted climb performance. These check climbs may be flown with the wings maintained in a near level attitude. Reciprocal climbs may not be necessary if inertial corrections (or another equivalent means) are applied to account for wind gradients. 4.9.2.2 If full rudder with wings level cannot maintain constant heading, small bank angles of up to 2° to 3° into the operating engine(s), with full rudder, should be used to maintain constant heading. Unless the landing lights automatically retract with engine failure, testing should be conducted with 4-68 ### Page 104 AC 25-7D CHG 1 the lights extended for § 25.121(a) Takeoff; landing gear extended, § 25.121(b) Takeoff; landing gear retracted, and § 25.121(d) Approach. 4.9.2.3 The climb performance tests with landing gear extended, in accordance with § 25.121(a), may be conducted with the landing gear and gear doors in a stable fully extended position. However, the critical configuration for the landing gear extended climb is considered to be that which presents the largest frontal area to the local airflow. This would normally be with no weight on the landing gear (full strut extension and trucks tilted) and all gear doors open. Since the takeoff path will be determined either from continuous takeoffs or checked by continuous takeoffs if constructed by the segmental method (Refer to § 25.111(d)), any non-conservatism arising from the gear doors “closed” climb data will be evident and should be corrected for. Also, some measure of conservatism is added to the landing gear extended climb performance by the requirement of § 25.111(d)(3) for the takeoff path data to be based on the airplane’s performance without ground effect. While during an actual takeoff the airplane may accelerate from VLOF towards V2, the climb gradient for showing compliance with § 25.121(a) is based on the VLOF speed as specified in the rule. 4.9.2.4 If means, such as variable intake doors, are provided to control powerplant cooling air supply during takeoff, climb, and en route flight, they should be set in a position that will maintain the temperature of major powerplant components, engine fluids, etc., within the established limits. The effect of these procedures should be included in the climb performance of the airplane. These provisions apply for all ambient temperatures up to the highest operational temperature limit for which approval is desired. (See § 25.1043.) 4.9.2.5 The latter parts of § 25.121(a)(1) and (b)(1), which state “...unless there is a more critical power operating condition existing later along the flight path...” are intended to cover those cases similar to where a wet engine depletes its water and reverts to dry engine operation. This is not intended to cover normal altitude power or thrust lapse rates above the point where retraction of the landing gear is begun. 4.9.2.6 Section 25.121(d) requires that the reference stall speed for the approach configuration not exceed 110 percent of the reference stall speed for the related landing configuration. This stall speed ratio requirement is to ensure that an adequate margin above the stall speed in the selected approach configuration is maintained during flap retraction in a go-around. An alternative means of providing an adequate operating speed margin during flap retraction in a go-around would be to increase VREF for the landing configuration to provide an equivalent operating speed margin. That is, VREF could be increased such that the reference stall speed for the 4-69 ### Page 105 **AC 25-7D CHG 1** approach configuration does not exceed 110 percent of VREF/1.23. An equivalent level of safety finding should be used to document the use of this alternative versus direct compliance with § 25.121(d). To maintain equivalent safety, the increase in VREF should not be excessive (for example, greater than 5 knots) to minimize the effect on safety of longer landing distances, higher brake energy demands, and reduced margins between VREF and VFE. 4.9.2.7 Section 25.121(d) permits the use of a climb speed established in connection with normal landing procedures, but not more than 1.4 VSR. Section 25.101(g) requires that the procedures for the execution of missed approaches associated with the conditions prescribed in § 25.121(d) be established. Consequently, the speeds and flap configuration used to show compliance with the minimum climb gradient requirements of § 25.121(d) need to be consistent with the speeds and flap configurations specified for go-around in the AFM operating procedures. To demonstrate the acceptability of recommended procedures, the applicant should conduct go-around demonstrations to include a weight, altitude, temperature (WAT)-limited or simulated WAT-limited thrust condition. In accordance with § 25.101(h), the established procedures must:  Be able to be consistently executed in service by crews of average skill,  Use methods or devices that are safe and reliable, and  Include allowance for any time delays in the execution of the procedures that may reasonably be expected in service. 4.9.2.8 FAA policy, as explained in policy PS-ANM100-1995-00058, Go-Around Power/Thrust Settings on Transport Category Airplanes, dated August 15, 1995, states that there should only be one power/thrust setting procedure used to show compliance with both §§ 25.119 and 25.121(d). 4.9.2.8.1 This policy is based on crew workload issues as discussed in the preamble for the ATTCS final rule (amendment 25-62): 4.9.2.8.2 That preamble states, “. . . current regulations preclude a higher thrust for approach climb (§ 25.121(d)) than for landing climb (§ 25.119). The workload required for the flightcrew to monitor and select from multiple inflight thrust settings in the event of an engine failure during a critical point in the approach, landing, or go-around operation is excessive.” 4.9.2.8.3 If the approach climb power/thrust setting is higher than the landing climb power/thrust setting, a throttle push would be required to obtain the AFM performance in the event of an engine failure after an all-engines-operating go-around has been initiated. The FAA considers the 4-70 ### Page 106 **AC 25-7D CHG 1** need to manually reset the engine power/thrust setting in a high workload environment to be unacceptable. 4.9.2.8.4 Systems that automatically reset power/thrust to a higher value after an engine failure (for example, ATTCS used for go-arounds) have been found acceptable as long as there is a single go-around power/thrust setting procedure for the one-engine-inoperative (approach climb) and all-engines-operating (landing climb) conditions. 4.10 En Route Flight Paths—§ 25.123. 4.10.1 Explanation. This guidance is intended for showing compliance with the requirements of § 25.123. 4.10.2 Procedures. 4.10.2.1 Sufficient en route climb performance data should be presented in the AFM to permit the determination of the net climb gradient and the net flight path in accordance with § 25.123(b) and (c) for all gross weights, altitudes, and ambient temperatures within the operating limits of the airplane. This en route climb performance data should be presented for altitudes up to the all-engines-operating ceiling to permit the calculation of drift-down data in the event of an en route engine failure. 4.10.2.2 Fuel Consumption Accountability. The effect of the variation of the airplane’s weight along the flight path due to the progressive consumption of fuel may be taken into account using fuel flow rates obtained from airplane manufacturers’ test data. If measured fuel flow data is unavailable, a conservative fuel flow rate not greater than 80 percent of the engine specification flow rate at maximum continuous thrust (MCT) may be used. 4.10.2.3 The procedures and flight conditions upon which the en route flight path data are based should be provided to the flightcrew. Credit for fuel dumping, if available and included in the flightcrew procedures, may be used to achieve the performance capability presented in the AFM. A conservative analysis should be used in taking into account the ambient conditions of temperature and wind existing along the flight path. All performance should be based on the net flight path and with MCT on the operating engine(s). 4-71 ### Page 107 **AC 25-7D CHG 1** 4.11 Landing—§ 25.125. 4.11.1 Explanation. 4.11.1.1 The landing distance is the horizontal distance from the point at which the main gear of the airplane is 50 feet above the landing surface (treated as a horizontal plane through the touchdown point) to the position of the nose gear when the airplane is brought to a stop. (For water landings, a speed of approximately 3 knots is considered “stopped.”) The beginning of the landing distance is referenced to the main gear because it is the lowest point of the airplane when the airplane is 50 feet above the landing surface. The end of the landing distance is referenced to the nose gear because it is the most forward part of the airplane in contact with the landing surface, and it should not extend beyond the certified landing distance. In this AC, the landing distance is divided into two parts: the airborne distance from 50 feet to touchdown, and the ground distance from touchdown to stop. The latter may be further subdivided into a transition phase and a full braking phase if the applicant prefers this method of analysis. 4.11.1.2 The minimum allowable value of VREF is specified in § 25.125(a)(2)(i) and (ii). It is intended to provide an adequate margin above the stall speed to allow for likely speed variations during an approach in light turbulence and to provide adequate maneuvering capability. If the landing demonstrations show that a higher speed is needed for acceptable airplane handling characteristics, the landing distance data presented in the AFM must be based upon the higher reference landing speed per § 25.125(b)(2). Further, if procedures recommend the use of approach speeds that are higher than VREF for reasons other than wind, flight tests should be conducted to determine whether the recommended VREF speeds are readily achievable at the landing threshold. If VREF is not readily achievable, then the AFM landing distances must include the effect of the excess speed at the landing threshold. 4.11.1.3 The engines should be set to the high side of the flight idle trim band, if applicable, for the landing flight tests. The effect of any variation in the idle fuel flow schedule for engines with electronic fuel controllers is typically negligible (but any such claim should be adequately substantiated). 4.11.2 Procedures for Determination of the Airborne Distance. Three acceptable means of compliance are described in paragraphs 4.11.2.1, 4.11.2.2 and 4.11.2.3 on the following pages. Note: If it is determined that the constraints on approach angle and touchdown rate-of-sink described in paragraphs 4.11.2.2 and 4.11.2.3 below are not appropriate due to novel or unusual features of the airplane’s design, new criteria may be 4-72 ### Page 108 AC 25-7D CHG 1 established. Such a change would be acceptable only if it is determined that an equivalent level of safety to existing performance standards and operational procedures is maintained. 4.11.2.1 Experience shows an upper bound to the part 25 zero-wind airborne distances achieved in past certifications and, similarly, a minimum speed loss. 4.11.2.1.1 These are approximated by the following: distance in lieu of measuring airborne distance and speed loss. If an applicant chooses to use these relationships, the applicant should show by test or analysis that they do not result in air distances or touchdown speeds that are nonconservative. 4.11.2.2 If an applicant chooses to measure airborne distance or time, at least six tests covering the landing weight range are required for each airplane configuration for which certification is desired. These tests should meet the following criteria: 4.11.2.2.1 A stabilized approach, targeting a glideslope of -3° and an indicated airspeed of VREF, should be maintained for a sufficient time prior to reaching a height of 50 feet above the landing surface to simulate a continuous approach at this speed. During this time, there should be no appreciable change in the power or thrust setting, pitch attitude, or rate of descent. The average glideslope of all landings used to show compliance should not be steeper than -3°. 4.11.2.2.2 Below 50 feet, there should be no nose depression by use of the longitudinal control and no change in configuration that requires action by the pilot, except for reduction in power or thrust. 4.11.2.2.3 The target rate of descent at touchdown should not exceed 6 feet per second. Although target values may not be precisely achieved, the average touchdown rate of descent should not exceed 6 feet per second. 4.11.2.3 If the applicant conducts enough tests to allow a parametric analysis (or equivalent method) that establishes, with sufficient confidence, the relationship between airborne distance (or time) as a function of the rates of descent at 50 feet and touchdown, the part 25 airborne distances may be based on an approach angle of -3.5°, and a touchdown sink rate of 8 feet per second. (See paragraph 4.11.8 for an example of this analysis method.) The parametric analysis method with these approach angle and touchdown 4-73 ### Page 109 AC 25-7D CHG 1 sink rate values should only be used for landing distances for which the operational safety margins required by § 121.195(b) or (c), § 135.385(b), (c), or (f), or equivalent will be applied. 4.11.2.3.1 At a given weight, the air distance or air time established by this method should not be less than 90 percent of the lowest demonstrated value obtained using the target values for approach angle and touchdown sink rate specified in paragraph 4.11.2.3.2 below. Test data with approach angles steeper than -3.5°, or touchdown sink rates greater than 8 feet per second, should not be used to satisfy this requirement. 4.11.2.3.2 In order to determine the parametric relationships, it is recommended that test targets span approach angles from -2.5° to –3.5°, and sink rates at touchdown from 2 to 6 feet per second. Target speed for all tests should be VREF. 4.11.2.3.3 Below 50 feet, there should be no nose depression by use of the longitudinal control and no change in configuration that requires action by the pilot, except for reduction in power or thrust. 4.11.2.3.4 If an acceptable method of analysis is developed by the applicant, a sufficient number of tests should be conducted in each aerodynamic configuration for which certification is desired to establish a satisfactory confidence level for the resulting air distance. Autolands may be included in the analysis but should not comprise more than half of the data points. If it is apparent that configuration is not a significant variable, all data may be included in a single parametric analysis. 4.11.2.3.5 If an applicant proposes any other method as being equivalent to a parametric analysis, that method should be based on a developed mathematical model that employs performance-related variables such as power or thrust, attitude, angle-of-attack, and load factor to adequately reproduce the flight test trajectory and airspeed variation from the 50-foot point to touchdown. Such a mathematical model should be validated by a sufficient number of tests to establish a satisfactory confidence level and be justified by a comparison of tested and calculated landing airborne distances. 4.11.2.3.6 For a derivative airplane with an aerodynamic configuration that has been previously certificated, if new tests are necessary to substantiate performance to a weight higher than that permitted by the extrapolation limits of § 25.21(d), two landings per configuration should be conducted for each 5 percent increase in landing weight (but no more than a total of six landings should be needed). These may be merged with previous certification tests for parametric analysis, regardless of whether the previous certification was conducted by this method or not. If a new 4-74 ### Page 110 **AC 25-7D CHG 1** aerodynamic configuration is proposed, the guidance described in paragraph 4.11.2.3.4 above, should be used. 4.11.2.3.7 In calculating the AFM landing distances, the speed loss from 50 feet to touchdown, as a percentage of VREF, may be determined using the conditions described in paragraph 4.11.2.3. 4.11.2.4 Whichever method is chosen to establish airborne distances, satisfactory flight characteristics should be demonstrated in the flare maneuver when a final approach speed of VREF-5 knots is maintained down to 50 feet. 4.11.2.4.1 Below 50 feet, the application of longitudinal control to initiate flare should occur at the same altitude as for a normal “on-speed” landing; no nose depression should be made and power or thrust should not be increased to facilitate the flare. 4.11.2.4.2 All power/thrust levers should be in their minimum flight idle position prior to touchdown. 4.11.2.4.3 The normal flare technique should be used, resulting in a touchdown speed approximately 5 knots less than the touchdown speed used to establish the landing distance. The rate of descent at touchdown should not be greater than 6 feet per second. 4.11.2.4.4 This demonstration should be performed over a range of weights (typically at maximum landing weight and near minimum landing weight), or at the most critical weight and CG combination as established by analysis or other acceptable means. 4.11.2.4.5 These VREF-5 knots landing demonstrations should not require the use of high control forces or full control deflections. 4.11.3 Procedures for Determination of the Transition and Stopping Distances. 4.11.3.1 The transition distance extends from the initial touchdown point to the point where all approved deceleration devices are operating. The stopping distance extends from the end of transition to the point where the airplane is stopped. The two phases may be combined at the applicant’s option. 4.11.3.2 If sufficient data are not available, there should be a minimum of six landings in the primary landing configuration. Experience has shown that if sufficient data are available for the airplane model to account for variation of braking performance with weight, lift, drag, ground speed, torque limit, etc., at least two test runs are necessary for each configuration when correlation for multiple configurations is being shown. 4-75 ### Page 111 **AC 25-7D CHG 1** 4.11.3.3 A series of at least six measured landing tests covering the landing weight range should be conducted on the same set of wheels, tires, and brakes in order to substantiate that excessive wear of wheel brakes and tires is not produced in accordance with the provisions of § 25.125(c)(2). The landing tests should be conducted with the normal operating brake pressures for which the applicant desires approval. The brakes may be in any wear state as long as an acceptable means is used to determine the landing distances with fully worn brakes for presentation in the AFM. The main gear tire pressure should be set to not less than the maximum pressure desired for certification corresponding to the specific test weight. Longitudinal control and brake application procedures should be such that they can be consistently applied in a manner that permits the airplane to be de-rotated at a controlled rate to preclude an excessive nose gear touchdown rate and so that the requirements of § 25.125(b)(4) and (5) are met. Nose gear touchdown rates in the certification landing tests should not be greater than eight feet per second. Certification practice has not allowed manually applied brakes before all main gear wheels are firmly on the ground. An automatic braking system can be armed before touchdown. 4.11.3.4 Describe the airplane operating procedures appropriate for determination of landing distance in the performance section of the AFM. 4.11.3.5 Propeller pitch position used in determining the normal all-engines-operating landing stopping distance should be established using the criteria of § 25.125(g) for those airplanes that may derive some deceleration benefit from operating engines. Section 25.125(g) states that if the landing distance determined using a “device” that depends on the operation of any engine would be “noticeably increased” when a landing is made with that engine inoperative, the landing distance must be determined with that engine inoperative, unless a “compensating means” will result in one-engine-inoperative landing distances not greater than those with all engines operating. Acceptable interpretations of the terms “device,” “noticeably increased,” and “compensating means” are described below. 4.11.3.5.1 If, with the normal operational ground idle setting procedure, the propeller produces drag at any speed during the stopping phase of the normal all-engines-operating landing distance, the maximum drag from this “device” for which performance credit may be taken is that which results from a propeller pitch position that gives not more than a slight negative thrust at zero airspeed. A slight negative thrust is that which will not cause the airplane, at light weight and without brakes being applied, to roll on a level surface. If the normal operational ground idle setting produces greater negative thrust at zero airspeed, the all-engines-operating stopping distances should be determined using a special flight test power lever stop to limit the propeller blade angle. 4-76 ### Page 112 **AC 25-7D CHG 1** 4.11.3.5.2 Distances should be measured for landings made with the propeller feathered on one engine, and ground idle selected after touchdown on the operating engines. The airplane configuration for this test, including the ground idle power lever position, should be the same as that used for the all-engines-operating landing distance determination. Differential braking may be used to maintain directional control. This testing should be conducted at the critical weight/CG position and landing speed. The propeller/engine rigging should be at the most adverse allowable tolerance. If the resulting distance does not exceed the all-engines-operating landing distance by more than two percent (2 percent), it is not “noticeably increased” and no further testing is required to take performance credit for all-engines-operating ground idle drag in the certified landing distances. 4.11.3.5.3 If the distances determined in paragraph 4.11.3.5.2 above are more than two percent greater than the all-engines-operating landing distances, there should be a “compensating means” in order to take performance credit for the all-engines-operating ground idle drag. Reverse propeller thrust on the operating engines is considered a “compensating means” if the resulting landing distances, with one propeller feathered, are demonstrated to be not longer than those determined for all-engines-operating with the ground idle setting. The airplane configuration for this test should be the same as that used for the all-engines-operating landing distance determination, except that the propeller reverse thrust position is used. The nose wheel should be free to caster, as in VMCG tests, to simulate wet runway surface conditions. Differential braking may be used to maintain directional control. Procedures for using propeller reverse thrust during the landing must be developed and demonstrated. The procedures associated with the use of propeller reverse thrust, required by § 25.101(f), must meet the requirements of § 25.101(h). The criteria outlined below may be applied to derive the levels of propeller reverse thrust consistent with recommended landing procedures and provide an acceptable means of demonstrating compliance with these requirements. This testing should be conducted at the critical weight/CG position and landing speed. The propeller/engine rigging should be at the most adverse allowable tolerance. If the “compensating means” do not allow performance credit for the all-engines-operating ground idle drag, a minimum of three weights that cover the expected range of operational landing weights and speeds should be tested. 4.11.3.5.4 In accordance with § 25.101(f), procedures for using propeller reverse thrust during landing must be developed and demonstrated. These procedures should include all of the pilot actions necessary to obtain the recommended level of propeller reverse thrust, maintain directional control, ensure safe engine operating characteristics and cancel propeller reverse thrust. 4-77 ### Page 113 **AC 25-7D CHG 1** 4.11.3.5.5 It should be demonstrated that using propeller reverse thrust during a landing complies with the engine operating characteristics requirements of § 25.939. The propeller reverse thrust procedures may specify a speed at which the propeller reverse thrust is cancelled in order to maintain safe engine operating characteristics. 4.11.3.5.6 The time sequence for the actions necessary to obtain the recommended level of propeller reverse thrust should be demonstrated by flight test. The time sequence used to determine the landing distances should reflect the most critical case relative to the time needed to obtain selected propeller reverse thrust. 4.11.3.5.7 The response times of the affected airplane systems to pilot inputs should be taken into account, for example, delays in system operation, such as interlocks and power lever detents that prevent the pilot from immediately selecting propeller reverse thrust. The effects of transient response characteristics, such as propeller reverse thrust engine spin-up, should also be included. 4.11.3.5.8 To enable a pilot of average skill to consistently obtain the recommended level of propeller reverse thrust under typical in-service conditions, a lever position that incorporates tactile feedback (e.g., a detent or stop) should be provided. If tactile feedback is not provided, a conservative level of propeller reverse thrust should be assumed. 4.11.3.5.9 The applicant should demonstrate that exceptional skill is not required to maintain directional control on a wet runway. The propeller reverse thrust procedures may specify a speed at which the propeller reverse thrust is cancelled in order to maintain directional controllability. 4.11.3.5.10 Compliance with the requirements of §§ 25.901(b)(2), 25.901(c), 25.1309(b), and 25.1309(c) will be accepted as providing compliance with the “safe and reliable” requirements of §§ 25.101(h)(2) and 25.125(c)(3). 4.11.4 Instrumentation and Data. Instrumentation should include a means to record the airplane’s glide path relative to the ground, and the ground roll against time, in a manner that permits determining the horizontal and vertical distance time-histories. The appropriate data to permit analysis of these time-histories should also be recorded. 4.11.5 Landing on Unpaved Runways. Guidance material for evaluation of landing on unpaved runways is contained in chapter 42 of this AC. 4.11.6 Automatic Braking Systems. 4-78 ### Page 114 **AC 25-7D CHG 1** Guidance material relative to evaluation of auto-brake systems is provided in paragraph 15.4.9 of this AC. 4.11.7 AFM Landing Distances. 4.11.7.1 In accordance with § 25.101(i), AFM landing distances must be determined with all the airplane wheel brake assemblies at the fully worn limit of their allowable wear range. The brakes may be in any wear state during the flight tests used to determine the landing distances, as long as a suitable combination of airplane and dynamometer tests is used to determine the landing distances corresponding to fully worn brakes. Alternatively, the relationship between brake wear and stopping performance established during accelerate-stop testing may be used if it encompasses the brake wear conditions and energies achieved during the airplane flight tests used to establish the landing distances. 4.11.7.2 In deriving the scheduled distances, the time delays shown in figure 4-19 below should be assumed. Figure 4-19. Landing Time Delays 4.11.7.2.1 Segment  represents the flight test measured average time from touchdown to pilot activation of the first deceleration device. For AFM data expansion, use the longer of 1 second or the test time. 4.11.7.2.2 Segment  represents the flight test measured average test time from pilot activation of the first deceleration device to pilot activation of the second deceleration device. For AFM data expansion, use the longer of 1 second or the test time. 4-79 ### Page 115 **AC 25-7D CHG 1** 4.11.7.2.3 Segment  is repeated until pilot activation of all deceleration devices has been completed and the airplane is in the full braking configuration. 4.11.7.3 For approved automatic deceleration devices (e.g., autobrakes or auto-spoilers, etc.) for which performance credit is sought for AFM data expansion, established times determined during certification testing may be used without the application of the 1-second minimum time delay required in the appropriate segment above. 4.11.7.4 It has been considered acceptable to expand the airborne portion of the landing distance in terms of a fixed airborne time, independent of airplane weight or approach speed. 4.11.7.5 Assumptions to be made in assessing the effect of wind on landing distance are discussed in paragraph 3.1 of this AC. 4.11.8 Parametric Analysis Data Reduction. The following is an acceptable method of converting the test data to a mathematical model for the parametric analysis method of air distance described in paragraph 4.11.2.3. 4.11.8.1 Test data for each test point: 4.11.8.2 The multiple linear regression analysis as outlined below is used to solve for the constants in the following equation: The form of the dependent variable being solved in the above equation is 4.11.8.3 50/t, rather than just t, in order to maintain the same units for all variables. 4-80 ### Page 116 **AC 25-7D CHG 1** 4.11.8.4 The test values of all the test points, 1 through n, are used to determine the constants a, b, and c in the above equation as follows, where n equals the number of test points and R1 through R13 are the regression coefficients: 4.11.8.5 Using the same regression coefficient relationships, determine the values of the constants, a, b, and c, for the speed reduction between 50 feet and 4-81 ### Page 117 **AC 25-7D CHG 1** touchdown (V50/VTD) by using the value of (V50/VTD) for (50/t) for each test point. 4.11.8.6 After determining the values of the constants, use the above equation for (50/t) to calculate the time from 50 feet to touchdown for the target conditions of a -3.5° flight path angle and R/STD = 8 ft/sec. Use a value of (R/S50) calculated from the approach path and V50. Then, using the same equation, but substituting (V50/VTD) for (50/t) and using the constants determined for (V50/VTD), calculate (V50/VTD). 4.11.8.7 After VTD is determined (from V50/VTD and V50), the air distance may be determined for the average flare speed and air time. Example Test Data: Run R/S 50 R/STD V50 VTD t 1 13.4 6.1 219 214 5.6 2 10.9 1.8 223 218 8.5 3 7.9 5.8 209 201 7.4 4 8.3 2.3 213 206 9.6 5 9.8 4.1 218 212 7.5 Results: 50/t = 1.0432 + 0.3647(R/S50) + 0.4917(R/STD) V50/VTD = 1.05508 - 0.003198(R/S50) + 0.001684(R/STD) For conditions of V50 = 220 ft/sec; flight path = -3.5°; R/STD = 8.0 ft/sec; the results are: R/S50 = 13.43 ft/sec V50/VTD = 1.0256 t = 5.063 sec Air distance = 1100 ft 4-82 ### Page 118 **AC 25-7D CHG 1** CHAPTER 5. FLIGHT: CONTROLLABILITY AND MANEUVERABILITY 5.1 General—§ 25.143. 5.1.1 Explanation. 5.1.1.1 The purpose of § 25.143 is to verify that any operational maneuvers conducted within the operational envelope can be accomplished smoothly with average piloting skill and without encountering a stall warning or other characteristics that might interfere with normal maneuvering, or without exceeding any airplane structural limits. Control forces should not be so high that the pilot cannot safely maneuver the airplane. Also, the forces should not be so light that it would take exceptional skill to maneuver the airplane without over-stressing it or losing control. The airplane response to any control input should be predictable to the pilot. 5.1.1.2 The maximum forces given in the table in § 25.143(d) for pitch and roll control for short term application are applicable to maneuvers in which the control force is only needed for a short period. Where the maneuver is such that the pilot will need to use one hand to operate other controls (such as during the landing flare or a go-around, or during changes of configuration or power/thrust resulting in a change of control force) the single-handed maximum control forces will be applicable. In other cases (such as takeoff rotation, or maneuvering during en route flight), the twohanded maximum forces will apply. 5.1.1.3 Short-term and long-term forces should be interpreted as follows: 5.1.1.3.1 Short-term forces are the initial stabilized control forces that result from maintaining the intended flight path following configuration changes and normal transitions from one flight condition to another, or from regaining control following a failure. By definition, short-term forces do not exist for any significant duration, and do not include momentary or transient peak forces that do not directly affect flight path, due to their short duration. 5.1.1.3.2 Long-term forces are those control forces that result from normal or failure conditions that cannot readily be trimmed out or eliminated. 5.1.1.4 In conducting the controllability and maneuverability tests to show compliance with § 25.143 at speeds between VMO/MMO and VFC/MFC, the airplane should be trimmed at VMO/MMO. 5.1.1.5 Modern wing designs can exhibit a significant reduction in maximum lift capability with increasing Mach number. The magnitude of this Mach number effect depends on the design characteristics of the particular wing. For wing designs with a large Mach number effect, the maximum bank 5-1 ### Page 119 **AC 25-7D CHG 1** angle that can be achieved while retaining an acceptable stall margin can be significantly reduced. Because the effect of Mach number can be significant, and because it can also vary greatly for different wing designs, the multiplying factors applied to VSR may be insufficient to ensure that adequate maneuvering capability exists at the minimum operating speeds. To address this issue, § 25.143(h) was added by amendment 25-108 to require a minimum bank angle capability in a coordinated turn without encountering stall warning or any other characteristic (including the envelope protection features of fly-by-wire flight control systems or automatic power or thrust increases) that might interfere with normal maneuvering. The maneuvering requirements consist of the minimum bank angle capability the FAA deems adequate for the specified regimes of flight combined with additional bank angle capability to provide a safety margin for various operational factors. These operational factors include both potential environmental conditions (e.g., turbulence, wind gusts) and an allowance for piloting imprecision (e.g., inadvertent overshoots). The FAA considers the automatic application of power or thrust by an envelope protection feature to be a feature that might interfere with normal maneuvering because it will result in a speed increase and flight path deviation, as well as potentially increasing crew workload due to the unexpected power or thrust increase. 5.1.2 General Test Requirements. 5.1.2.1 Compliance with § 25.143 (a) through (g) is primarily a qualitative determination by the pilot during the course of the flight test program. The control forces required and airplane response should be evaluated during changes from one flight condition to another and during maneuvering flight. The forces required should be appropriate to the flight condition being evaluated. For example, during an approach for landing, the forces should be light and the airplane responsive in order that adjustments in the flight path can be accomplished with a minimum of workload. In cruise flight, forces and airplane response should be such that inadvertent control input does not result in exceeding limits or in undesirable maneuvers. Longitudinal control forces should be evaluated during accelerated flight to ensure a positive stick force with increasing normal acceleration. Forces should be heavy enough at the limit load factor to prevent inadvertent excursions beyond the design limit. Sudden engine failures should be investigated during any flight condition or in any configuration considered critical, if not covered by another section of part 25. Control forces considered excessive should be measured to verify compliance with the maximum control force limits specified in § 25.143(d). Allowance should be made for delays in the initiation of recovery action appropriate to the situation. 5-2 ### Page 120 5.1.2.2 Since § 25.143(h) involves a target speed, bank angle, and maximum value of thrust/power setting, not all flight test conditions to demonstrate compliance will necessarily result in a constant-altitude, thrust-limited turn. In cases with positive excess power or thrust, a climbing condition at the target bank and speed is acceptable. Alternately, if desired, the power or thrust may be reduced to less than the maximum allowed, so that compliance is shown with a completely stabilized, constant-altitude turn. With the airplane stabilized in a coordinated turn, holding power or thrust and speed, increase bank angle at constant airspeed until compliance is shown. For cases with negative excess power or thrust (e.g., the landing configuration case), a constant-altitude slow-down maneuver at the target bank angle has been shown to be a suitable technique. With the airplane descending at VREF in wings-level flight on a 3° glide path, trim and throttle position are noted. The airplane is then accelerated to VREF + 10 to 20 knots in level flight. The original trim and throttle conditions are reset as the airplane is rolled into a constant-altitude slow-down turn at the target bank angle. Throttles can be manipulated between idle and the marked position to vary slow-down rate as desired. Compliance is shown when the airplane decelerates through VREF in the turn without encountering a stall warning or other characteristic that might interfere with normal maneuvering. 5.1.2.3 If stall warning is provided by an artificial stall warning system, the effect of production tolerances on the stall warning system should be considered when evaluating compliance with the maneuvering capability requirements of § 25.143(h). See paragraph 8.1.6.2.6 of this AC for more information. 5.1.3 Controllability Following Engine Failure. Section 25.143(b)(1) requires the airplane to be controllable following the sudden failure of the critical engine. To show compliance with this requirement, the demonstrations described in paragraphs 5.1.3.1 and 5.1.3.2 below, should be made with engine failure (simulated by fuel cuts) occurring during straight, wings level flight. To allow for likely in-service delays in initiating recovery action, no action should be taken to recover control for two seconds following pilot recognition of engine failure. The recovery action should not necessitate movement of the engine, propeller, or trim controls, and should not result in excessive control forces. Additionally, the airplane will be considered to have reached an unacceptable attitude if the bank angle exceeds 45° during the recovery. These tests may be conducted using throttle slams to idle, with actual fuel cuts repeated only for those tests found to be critical. 5.1.3.1 At each takeoff flap setting at the initial all-engine climb speed (e.g., V2 + 10 knots) with: 5.1.3.1.1 All engines operating at maximum takeoff power or thrust prior to failure of the critical engine; 5-3 ### Page 121 5.1.3.1.2 All propeller controls (if applicable) in the takeoff position; 5.1.3.1.3 The landing gear retracted; and 5.1.3.1.4 The airplane trimmed at the prescribed initial flight condition. 5.1.3.2 With the wing flaps retracted at a speed of 1.23 VSR with: 5.1.3.2.1 All engines operating at maximum continuous power or thrust prior to failure of the critical engine; 5.1.3.2.2 All propeller controls in the en route position; 5.1.3.2.3 The landing gear retracted; and 5.1.3.2.4 The airplane trimmed at the prescribed initial flight condition. 5.1.4 Pilot Induced Oscillations (PIO). 5.1.4.1 Explanation. 5.1.4.1.1 Section 25.143(a) and (b) require that the airplane be safely controllable and maneuverable without exceptional piloting skill and without danger of exceeding the airplane limiting load factor under any probable operating conditions. Service history events have indicated that modern transport category airplanes can be susceptible to airplane-pilot coupling under certain operating conditions and would not meet the intent of this requirement. 5.1.4.1.2 The classic PIO is considered to occur when an airplane’s response is approximately 180° out of phase with the pilot’s control input. However, PIO events with 180° phase relationships are not the only conditions in which the airplane may exhibit closed-loop (pilot-in-the-loop) characteristics that are unacceptable for operation within the normal, operational, or limit flight envelopes. Others include unpredictability of the airplane’s response to the pilot’s control input. This may be due to nonlinearities in the control system, actuator rate or position limiting not sensed by the pilot through the flight controls or changing pitch response at high altitude as the airplane maneuvers into and out of Mach buffet. Artificial trim and feel systems which produce controllers with too small a displacement and light force gradients may also lead to severe over control. This is especially true in a dynamic environment of high-altitude turbulence or upsets in which the autopilot disconnects. This places the airplane in the hands of the unsuspecting pilot in conditions of only a small g or airspeed margin to buffet onset and with very low aerodynamic damping. These characteristics, while not a classic 180° out of phase PIO per se, may be hazardous and should be considered under the more general description of airplane-pilot coupling tendencies. 5-4 ### Page 122 **AC 25-7D CHG 1** 5.1.4.1.3 Some of the PIO tendency characteristics described in paragraph 5.1.4.1.2 above are attributes of transport airplanes (e.g., low frequency short period, large response lags) that are recognized by part 25. Limits are placed on some of these individual attributes by part 25 (e.g., stick force per g, heavily damped short period) to assure satisfactory open-loop characteristics. However, service reports from recent years have indicated that certain operating envelope conditions, combined with triggering events, can result in airplane-pilot coupling incidents. Some of the conditions that have led to these PIOs include fuel management systems that permit extended operations with a CG at or near the aft limit, operating at weight/speed/altitude conditions that result in reduced margins to buffet onset combined with tracking tasks such as not exceeding speed limitations and severe buffet due to load factor following an upset, and control surface rate or position limiting. 5.1.4.1.4 This service experience has shown that compliance with only the quantitative, open-loop (pilot-out-of-the loop) requirements does not guarantee that the required levels of flying qualities are achieved. Therefore, to ensure that the airplane has achieved the flying qualities required by § 25.143(a) and (b), the airplane should be evaluated by test pilots conducting high-gain (wide-bandwidth), closed-loop tasks to determine that the potential of encountering adverse PIO tendencies is minimal. 5.1.4.1.5 For the most part, these tasks should be performed in actual flight. However, for conditions that are considered too dangerous to attempt in actual flight (i.e., certain flight conditions outside of the operational flight envelope, flight in severe atmospheric disturbances, flight with certain failure states, etc.), the closed loop evaluation tasks may be performed using a motion base high fidelity simulator if it can be validated for the flight conditions of interest. 5.1.4.2 Special Considerations. 5.1.4.2.1 The certification team should understand the flight control system and airplane design. 5.1.4.2.2 The applicant should explain why the design is not conducive to a PIO problem and how this is to be shown in both developmental and certification flight tests. 5.1.4.2.3 The applicant should explain what has been done during the development flight test experience and any design changes that were required for PIO problems. 5.1.4.2.4 The certification flight test program should be tailored to the specific airplane design and to evaluate the airplane in conditions that were found 5-5 ### Page 123 AC 25-7D CHG 1 to be critical during its development program and PIO analytical assessment. 5.1.4.2.5 The FAA flight test pilots should also continuously evaluate the airplane for PIO tendencies during the certification program in both the airplane and simulator. This evaluation should include both normal and malfunction states; all certification flight test points; transitions between and recoveries from these flight test points; and normal, crosswind, and offset landing task evaluations. 5.1.4.2.6 Since the evaluation of flying qualities under § 25.143(a) and (b) is basically qualitative, especially evaluations of PIO susceptibility, the high-gain tasks discussed herein should be accomplished by at least three test pilots. Use of other pilots can provide additional insights into the airplane handling qualities, but for the purpose of demonstrating compliance with this requirement the evaluation pilots should be trained test pilots. 5.1.4.3 Procedures: Flight Test. 5.1.4.3.1 Evaluation of the actual task performance achieved, e.g., flight technical error, is not recommended as a measure of proof of compliance. Only the pilot’s rating of the PIO characteristics is needed as described in paragraph 5.1.4.6. The tasks are used only to increase the pilot’s gain, which is a prerequisite for exposing PIO tendencies. Although task performance is not used as proof of compliance, task performance should be recorded and analyzed to insure that all pilots seem to be attempting to achieve the same level of performance. 5.1.4.3.2 Tasks for a specific certification project should be based on operational situations, flight testing maneuvers, or service difficulties that have produced PIO events. Task requirements for a specific project will be dictated by the particular airplane and its specific areas of interest as determined by the tailored flight test program mentioned above. Some of these include high altitude upset maneuvers, encounters with turbulence at high altitude in which the autopilot disconnects, crosswind/crossed control landings with and without one engine inoperative, and offset landings to simulate the operational case in which the airplane breaks out of instrument meteorological conditions (IMC) offset from the glideslope and/or localizer beam and the pilot makes a rapid alignment correction. Tests should be conducted at or near the critical altitude/weight/CG combinations. 5.1.4.3.3 Tasks described here may be useful in any given evaluation and have proven to be operationally significant in the past. It is not intended that these are the only tasks that may be used or may be required depending on the scope and focus of the individual evaluation being conducted. Other 5-6 ### Page 124 AC 25-7D CHG 1 tasks may be developed and used as appropriate. For example, some manufacturers have used formation tracking tasks successfully in the investigation of these tendencies. For all selected tasks, a build-up approach should be used, and all end points should be approached with caution. Capture tasks and fine tracking tasks share many common characteristics but serve to highlight different aspects of any PIO problem areas that may exist. In some cases, depending on individual airplane characteristics, it may be prudent to look at capture tasks first and then proceed to fine tracking tasks or combined gross acquisition (capture) and fine tracking tasks as appropriate. 5.1.4.4 Capture Tasks. 5.1.4.4.1 Capture tasks are intended to evaluate handling qualities for gross acquisition as opposed to continuous tracking. A wide variety of captures can be done provided the necessary cues are available to the pilot. Pitch attitude, bank angle, heading, flight path angle, angle-of-attack, and g captures can be done to evaluate different aspects of the airplane response. These capture tasks can give the pilot a general impression of the handling qualities of the airplane, but because they do not involve closed-loop fine tracking, they do not expose all of the problems that may arise in fine tracking tasks. Capture tasks should not be used as the only evaluation tasks. 5.1.4.4.2 For pitch captures, the airplane is trimmed for a specified flight condition. The pilot aggressively captures 5° pitch attitude (or 10° if the airplane is already trimmed above 5°). The pilot then makes a series of aggressive pitch captures of 5° increments in both directions, and then continues this procedure with 10° increments in both directions. An airplane with more capability can continue the procedure with larger pitch excursions. If possible, the initial conditions for each maneuver should be such that the airplane will remain within ±1,000 feet and ±10 knots of the specified flight condition during the maneuver; however, large angle captures at high-speed conditions will inevitably produce larger speed and altitude changes. If the airplane should get too far from the specified condition during a task, it should be re-trimmed for the specified condition before starting the next maneuver. 5.1.4.4.3 The other kinds of captures are usually done in a similar manner, with some minor differences. G captures can be done from a constant-g turn or pull ups and pushovers using ±0.2 g and ±0.5 g. Heading captures can be used to evaluate the yaw controller alone (usually small heading changes of 5° or less). 5.1.4.4.4 Bank angle captures are also commonly done using bank-to-bank rolls. Starting from a 15° bank angle, the pilot aggressively rolls and captures the opposite 15° bank angle (total bank angle change of 30°). The pilot 5-7 ### Page 125 AC 25-7D CHG 1 then rolls back and captures 15° bank in the original direction. This procedure should continue for a few cycles. The procedure is then repeated using 30° bank angles, and then repeated again using 45° bank angles. A variation of this is to capture wings-level from the initial bank condition. 5.1.4.4.5 Where suitable, combined conditions could be used as described in the task shown in paragraph 5.1.4.4.6 below, in which a target g and bank angle are tightly tracked until the target pitch attitude and heading are captured. 5.1.4.4.6 The following upset and/or collision avoidance maneuvers have been found to be effective in evaluating PIO susceptibility when the airplane is flying at high altitude under conditions of low g to buffet onset, typically 0.3 g. This emphasis on cruise susceptibility stems from operational experiences, but should not be interpreted as placing less emphasis on other flight phases. 1. Trim for level f light at long range cruise Mach number. Initiate a slight climb and slow the aircraft while leaving power/thrust set. Push the nose over and set up a descending turn with 30° to 40° of bank and approximately 10° nose below the horizon, or as appropriate, to accelerate to the initial trim speed. At the initial trim airspeed initiate a 1.5 g to 1.67 g (not to exceed deterrent buffet) pull up and establish a turn in the opposite direction to a heading which will intercept the initial course on which the airplane was trimmed. Establish a pitch attitude which will provide a stabilized climb back to the initial trim altitude. The pilot may use the throttles as desired during this maneuver and should pick a target g, bank angle, heading, and pitch attitude to be used prior to starting the maneuver. The target g and bank angle should be set and tightly tracked until the target pitch attitude and heading are obtained respectively. The stabilized steady heading climb should be tightly tracked for an adequate amount of time to allow the pilot to assess handling qualities, even through the initial trim altitude and course if required. The pilot should qualitatively evaluate the airplane during both the gross acquisition and fine tracking portions of this task while looking for any tendency towards PIO in accordance with the criteria in paragraph 5.1.4.6. 2. This maneuver should be repeated in the nose-down direction by accelerating to MMO from the trim condition 10° nose down and then recover as above. 3. Trim for level flight as above. Initiate a 1.5 g to 1.67 g (not to exceed deterrent buffet) pull-up and approximately a 30° bank turn. Once the target g is set, transition the aircraft to approximately a 0.5 g pushover and reverse the turn to establish an intercept heading to the initial course. Using power or thrust as required, set up a stabilized steady 5-8 ### Page 126 AC 25-7D CHG 1 heading descent to intercept the initial course and altitude used for the trimmed condition. The pilot may continue the heading and descent through the initial conditions to allow more tracking time if needed. Attempt to precisely set and track bank angle, g, heading, and pitch attitude as appropriate. The pilot should qualitatively evaluate the airplane during both the gross acquisition and fine tracking portions of this task while looking for a PIO tendency in accordance with paragraph 5.1.4.6. 5.1.4.5 Fine Tracking Tasks. 5.1.4.5.1 These tasks may be used to assess the airplane’s PIO susceptibility when flying in turbulent atmospheric conditions. In this task, a tracking target is displayed which commands pitch and roll changes for the evaluation pilot to follow. Whatever visual cue is used (e.g., head up display (HUD), flight director, etc.), it should present the tracking task without filtering, smoothing, or bias. The pitch and roll commands should be combinations of steps and ramps. The sequence of pitch and roll commands should be designed to keep the airplane within ±1,000 feet of the test altitude and within ±10 knots of the test airspeed. The sequence should be long enough and complex enough that the pilot cannot learn to anticipate the commands. The unfamiliarity is intended to help keep the test pilot’s gain high and to preclude inadvertent pilot compensation while accomplishing the task. Such compensation, along with reduced gains, could mask any PIO tendencies. 5.1.4.5.2 Even though these fine tracking tasks will provide insight into PIO susceptibility of a conventional airplane when flying in turbulence, other considerations apply to augmented airplane types. For example, structural load alleviation systems that use the same flight control surface as the pilot will limit the pilot’s control authority in turbulent atmospheric conditions. Under these circumstances of rate or position limiting, PIO tendencies will be more critical as previously discussed. Therefore, specific evaluations for turbulent atmospheric conditions with these systems operating are necessary for these airplane types. 5.1.4.5.3 For single axis tasks, it has been found that aural commands given in a timed sequence provide an adequate cue in the event it is not possible to modify the flight director to display the pitch commands. 5.1.4.5.4 Based on PIO events seen in service, high altitude tracking tasks (with up to approximately ±4° pitch excursions from trim occurring at varying intervals of approximately 2 to 5 seconds) have been effective in evaluating PIO susceptibility. These tasks have been used where the 5-9 ### Page 127 AC 25-7D CHG 1 airplane is flying under conditions of low g margin to buffet onset. The time history in figure 5-1 below is a pictorial representation of a sample task in MIL-STD-1797A that has the desired attributes for high altitude PIO evaluations. Figure 5-1. Sample Pitch Tracking Task time (seconds) Δ pitch angle (degrees) -4 -3 -2 -1 0 1 2 3 4 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 5.1.4.6 PIO Assessment Criteria. 5.1.4.6.1 The evaluation of an airplane for PIO susceptibility will be conducted using the FAA handling qualities rating method (HQRM). (See appendix E of this AC for more information on the HQRM). Tasks should be designed to focus on any PIO tendencies that may exist. Table 5-1 below contains the descriptive material associated with PIO characteristics and its relationship to the PIO Rating Scale called out in the U.S. Military Standard. 5-10 ### Page 128 Table 5-1. PIO Rating Criteria and Comparison to MIL Standard FAA HQ Rating PIO Characteristics Description MIL-STD- 1797A PIO Rating Scale No tendency for pilot to induce undesirable motion. 1 SAT Undesirable motions (overshoots) tend to occur when pilot initiates abrupt maneuvers or attempts tight control. These motions can be prevented or eliminated by pilot technique. (No more than minimal pilot compensation is required.) 2 ADQ Undesirable motions (unpredictability or over control) easily induced when pilot initiates abrupt maneuvers or attempts tight control. These motions can be prevented or eliminated but only at sacrifice to task performance or through considerable pilot attention and effort. (No more than extensive pilot compensation is required.) 3 CON Oscillations tend to develop when pilot initiates abrupt maneuvers or attempts tight control. Adequate performance is not attainable and pilot has to reduce gain to recover. (Pilot can recover by merely reducing gain.) 4 UNSAT Divergent oscillations tend to develop when pilot initiates abrupt maneuvers or attempts tight control. Pilot has to open control loop by releasing or freezing the controller. 5 Disturbance or normal pilot control may cause divergent oscillation. Pilot has to open control loop by releasing or freezing the controller. 6 SAT = Satisfactory CON = Controllable ADQ = Adequate UNSAT = Unsatisfactory or Failed 5.1.4.6.2 Table 5-1 above provides the FAA handling qualities (HQ) rating descriptions of airplane motions that may be seen during the conduct of specific PIO tasks or during tests throughout the entire certification flight test program. The italicized phrases highlight major differences between rating categories in the table. 5.1.4.6.3 The acceptable HQ ratings for PIO tendencies is shown in table E-2 of appendix E. As described in that appendix, the minimum HQ rating, and consequently the pass/fail criteria, varies with the flight envelope, atmospheric disturbance considered, and failure state. For example, table 5-2 below shows a handling qualities matrix for a tracking task with 5-11 ### Page 129 the airplane at aft CG trimmed in flight conditions giving 1.3 g to buffet onset. Table 5-2. Example of Acceptable HQ Rating for PIO Tendencies Airspeed MLRC MLRC MLRC MLRC Load Factor Range 0.8 to 1.3 -1.0 to 2.5 0.8 to 1.3 -1.0 to 2.5 Buffet Level Onset Deterrent Onset Deterrent Turbulence Light Light Light Light Failure None None Improbable failure of SAS Improbable failure of SAS Flight Envelope NFE LFE NFE LFE Minimum Permitted HQ Rating SAT ADQ ADQ CON SAT = Satisfactory ADQ = Adequate CON = Controllable NFE = Normal flight envelope LFE = Limit flight envelope SAS = Stability augmentation system MLRC = Long range cruise Mach number 5.1.5 Maneuvering Characteristics—§ 25.143(g). 5.1.5.1 General. An acceptable means of compliance with the requirement that stick forces may not be excessive when maneuvering the airplane is to demonstrate that, in a turn for 0.5 g incremental normal acceleration (0.3 g above 20,000 ft) at speeds up to VFC/MFC, the average stick force gradient does not exceed 120 lbs per g. 5.1.5.2 Interpretive Material. The objective of § 25.143(g) is to ensure that the limit strength of any critical component on the airplane would not be exceeded in maneuvering flight. In much of the structure, the load sustained in maneuvering flight can be assumed to be directly proportional to the load factor applied. However, this may not be the case for some parts of the structure (e.g., the tail and rear fuselage). Nevertheless, it is accepted that the airplane load factor will be a sufficient guide to the possibility of exceeding limit 5-12 ### Page 130 strength on any critical component if a structural investigation is undertaken whenever the design positive limit maneuvering load factor is closely approached. If flight testing indicates that the positive design limit maneuvering load factor could be exceeded in steady maneuvering flight with a 50-lb stick force, the airplane structure should be evaluated for the anticipated load at a 50-lb stick force. The airplane will be considered to have been overstressed if limit strength has been exceeded in any critical component. For the purposes of this evaluation, limit strength is defined as the lesser of either the limit design loads envelope increased by the available margins of safety, or the ultimate static test strength divided by 1.5. 5.1.5.3 Minimum Stick Force to Reach Limit Strength. 5.1.5.3.1 A stick force of at least 50 lbs to reach limit strength in steady maneuvers or wind-up turns is considered acceptable to demonstrate adequate minimum force at limit strength in the absence of deterrent buffeting. If heavy buffeting occurs before the limit strength condition is reached, a somewhat lower stick force at limit strength may be acceptable. The acceptability of a stick force of less than 50 lbs at the limit strength condition will depend upon the intensity of the buffet, the adequacy of the warning margin (i.e., the load factor increment between the heavy buffet and the limit strength condition), and the stick force characteristics. In determining the limit strength condition for each critical component, the contribution of buffet loads to the overall maneuvering loads should be taken into account. 5.1.5.3.2 This minimum stick force applies in the en route configuration with the airplane trimmed for straight flight, at all speeds above the minimum speed at which the limit strength condition can be achieved without stalling. No minimum stick force is specified for other configurations, but the requirements of § 25.143(g) are applicable in these conditions. 5.1.5.4 Stick Force Characteristics. 5.1.5.4.1 At all points within the buffet onset boundary determined in accordance with § 25.251(e), but not including speeds above VFC/MFC, the stick force should increase progressively with increasing load factor. Any reduction in stick force gradient with change of load factor should not be so large or abrupt as to significantly impair the ability of the pilot to maintain control over the load factor and pitch attitude of the airplane. 5.1.5.4.2 Beyond the buffet onset boundary, hazardous stick force characteristics should not be encountered within the permitted maneuvering envelope as limited by paragraph 5.1.5.4.3. It should be possible, by use of the primary longitudinal control alone, to rapidly pitch the airplane nose down so as to 5-13 ### Page 131 regain the initial trimmed conditions. The stick force characteristics demonstrated should comply with the following: 1. For normal acceleration increments of up to 0.3 g beyond buffet onset, where these can be achieved, local reversal of the stick force gradient may be acceptable, provided that any tendency to pitch up is mild and easily controllable. 2. For normal acceleration increments of more than 0.3 g beyond buffet onset, where these can be achieved, more marked reversals of the stick force gradient may be acceptable. It should be possible to contain any pitch-up tendency of the airplane within the allowable maneuvering limits, without applying push forces to the control column and without making a large and rapid forward movement of the control column. 5.1.5.4.3 In flight tests to satisfy paragraphs 5.1.5.4.1 and 5.1.5.4.2 above, the load factor should be increased until: 1. The level of buffet becomes sufficient to provide a strong and effective deterrent to any further increase of the load factor; 2. Further increase of the load factor requires a stick force in excess of 150 lbs (or in excess of 100 lbs when beyond the buffet onset boundary) or is impossible because of the limitations of the control system; or 3. The positive limit maneuvering load factor established in compliance with § 25.337(b) is achieved. 5.1.5.5 Negative Load Factors. It is not intended that a detailed flight test assessment of the maneuvering characteristics under negative load factors should necessarily be made throughout the specified range of conditions. An assessment of the characteristics in the normal flight envelope involving normal accelerations from 1 g to zero g will normally be sufficient. Stick forces should also be assessed during other required flight testing involving negative load factors. Where these assessments reveal stick force gradients that are unusually low, or that are subject to significant variation, a more detailed assessment, in the most critical of the specified conditions, will be required. This may be based on calculations, provided they are supported by adequate flight test or wind tunnel data. 5.1.6 Thrust or Power Setting for Maneuver Capability Demonstrations. The effect of thrust or power on maneuver capability is normally a function of only the thrust-to-weight ratio. Therefore, for those configurations in which the WAT-limited thrust or power setting is prescribed, it is usually acceptable to use the thrust or power setting that is consistent with a WAT-limited climb gradient at the test conditions of weight, altitude, and temperature. However, if the maneuver margin to stall warning (or other characteristic that might interfere with normal maneuvering) is reduced with 5-14 ### Page 132 **AC 25-7D CHG 1** increasing thrust or power, the critical conditions of both thrust or power and thrust-to-weight ratio should be taken into account when demonstrating the required maneuvering capabilities. 5.2 Longitudinal Control—§ 25.145. 5.2.1 Explanation. 5.2.1.1 Section 25.145(a) requires that there be adequate longitudinal control to promptly pitch the airplane nose down from at or near the stall to return to the original trim speed. The intent is to ensure that there is sufficient pitch control for a prompt recovery if inadvertently slowed to the point of stall. Although this requirement must be met with power off and at maximum continuous thrust or power, there is no intention to require stall demonstrations with thrust or power above that specified in § 25.201(a)(2). Instead of performing a full stall at maximum continuous power or thrust, compliance may be assessed by demonstrating sufficient static longitudinal stability and nose down control margin when the deceleration is ended at least one second past stall warning during a one knot per second deceleration. The static longitudinal stability during the maneuver and the nose down control power remaining at the end of the maneuver must be sufficient to assure compliance with the requirement. 5.2.1.2 Section 25.145(b) requires changes to be made in flap position, power or thrust, and speed without undue effort when re-trimming is impractical. The purpose is to ensure that any of these changes are possible assuming that the pilot finds it necessary to devote at least one hand to the initiation of the desired operation without being overpowered by the primary airplane controls. The objective is to show that an excessive change in trim does not result from the application or removal of power or thrust or the extension or retraction of wing flaps. The presence of gated positions on the flap control does not affect the requirement to demonstrate full flap extensions and retractions without changing the trim control. Compliance with § 25.145(b) also requires that the relation of control force to speed be such that reasonable changes in speed may be made without encountering very high control forces. 5.2.1.3 Section 25.145(c) contains requirements associated primarily with attempting a go-around maneuver from the landing configuration. Retraction of the high-lift devices from the landing configuration should not result in a loss of altitude if the power or thrust controls are moved to the go-around setting at the same time that flap/slat retraction is begun. The design features involved with this requirement are the rate of flap/slat retraction, the presence of any flap gates, and the go-around power or thrust setting. The go-around power or thrust setting should be the same as is used to comply with the approach and landing climb performance 5-15 ### Page 133 AC 25-7D CHG 1 requirements of §§ 25.121(d) and 25.119, and the controllability requirements of §§ 25.145(b)(3), 25.145(b)(4), 25.145(b)(5), 25.149(f), and 25.149(g). The controllability requirements may limit the go-around power or thrust setting. 5.2.1.4 Section 25.145(d) provides requirements for demonstrating compliance with § 25.145(c) when gates are installed on the flap selector. Section 25.145(d) also specifies gate design requirements. Flap gates, which prevent the pilot from moving the flap selector through the gated position without a separate and distinct movement of the selector, allow compliance with these requirements to be demonstrated in segments. High lift device retraction must be demonstrated beginning from the maximum landing position to the first gated position, between gated positions, and from the last gated position to the fully retracted position. 5.2.1.4.1 If gates are provided, § 25.145(d) requires the first gate from the maximum landing position to be to be located at a position corresponding to a go-around configuration. If there are multiple go-around configurations, the following criteria should be considered when selecting the location of the gate: 1. The expected relative frequency of use of the available go-around configurations. 2. The effects of selecting the incorrect high-lift device control position. 3. The potential for the pilot to select the incorrect control position, considering the likely situations for use of the different go-around positions. 4. The extent to which the gate(s) aid the pilot in quickly and accurately selecting the correct position of the high-lift devices. 5.2.1.4.2 Regardless of the location of any gates, initiating a go-around from any of the approved landing positions should not result in a loss of altitude. Therefore, § 25.145(d) requires that compliance with § 25.145(c) be demonstrated for retraction of the high-lift devices from each approved landing position to the control position(s) associated with the high-lift device configuration(s) used to establish the go-around procedure(s) from that landing position. A separate demonstration of compliance with this requirement should only be necessary if there is a gate between an approved landing position and its associated go-around position(s). If there is more than one associated go-around position, conducting this test using the go-around configuration with the most retracted high-lift device position should suffice, unless there is a more critical case. If there are no gates between any of the landing flap positions and their associated go-around positions, the demonstrations discussed in paragraph 5.2.1.4 5-16 ### Page 134 **AC 25-7D CHG 1** above should be sufficient to show compliance with this provision of § 25.145(d). 5.2.2 Procedures. The following test procedures outline an acceptable means for demonstrating compliance with § 25.145. These tests may be conducted at an optional altitude in accordance with § 25.21(c). Where applicable, the conditions should be maintained on the engines throughout the maneuver. 5.2.2.1 Longitudinal Control Recovery—§ 25.145(a). 5.2.2.1.1 Configuration.  Maximum weight, or a lighter weight if more critical.  Critical CG position.  Landing gear extended.  Wing flaps retracted and extended to the maximum landing position.  Engine power or thrust at idle and maximum continuous. 5.2.2.1.2 Test Procedure. The airplane must be trimmed at the speed for each configuration as prescribed in § 25.103(b)(6). The airplane should then be decelerated at 1 knot per second with wings level. For tests at idle power or thrust, the applicant must demonstrate that the nose can be pitched down from any speed between the trim speed and the stall. Typically, the most critical point is at the stall when in stall buffet. The rate of speed increase during the recovery should be adequate to promptly return to the trim point. Data from the stall characteristics testing can be used to evaluate this capability at the stall. For tests at maximum continuous power or thrust, the maneuver need not be continued for more than one second beyond the onset of stall warning. However, the static longitudinal stability characteristics during the maneuver, and the nose down control power remaining at the end of the maneuver, must be sufficient to assure that a prompt recovery to the trim speed could be attained if the airplane is slowed to the point of stall. 5.2.2.2 Longitudinal Control, Flap Extension—§ 25.145(b)(1). 5.2.2.2.1 Configuration.  Maximum landing weight or a lighter weight if considered more critical.  Critical CG position.  Wing flaps retracted. 5-17 ### Page 135 **AC 25-7D CHG 1**  Landing gear extended.  Engine power or thrust at flight idle. 5.2.2.2.2 Test procedure. The airplane must be trimmed at a speed of 1.3 VSR. The flaps must be extended to the maximum landing position as rapidly as possible while maintaining approximately 1.3 VSR for the flap position existing at each instant throughout the maneuver. The control forces must not exceed 50 lbs (the maximum force for short term application that can be applied readily by one hand) throughout the maneuver without changing the trim control. 5.2.2.3 Longitudinal Control, Flap Retraction—§ 25.145(b)(2) and (3). 5.2.2.3.1 Configuration.  Maximum landing weight or a lighter weight if considered more critical.  Critical CG position.  Wing flaps extended to the maximum landing position.  Landing gear extended.  Engine power or thrust at flight idle and the go-around power or thrust setting. 5.2.2.3.2 Test procedure. With the airplane trimmed at 1.3 VSR, the flaps must be retracted to the full up position while maintaining approximately 1.3 VSR for the flap position existing at each instant throughout the maneuver. The longitudinal control force must not exceed 50 lbs throughout the maneuver without changing the trim control. 5.2.2.4 Longitudinal Control, Power or Thrust Application—§ 25.145(b)(4) and (5). 5.2.2.4.1 Configuration.  Maximum landing weight or a lighter weight if considered more critical.  Critical CG position.  Wing flaps retracted and extended to the maximum landing position.  Landing gear extended.  Engine power or thrust at flight idle. 5-18 ### Page 136 **AC 25-7D CHG 1** 5.2.2.4.2 Test procedure. The airplane must be trimmed at a speed of 1.3 VSR. Quickly set goaround power or thrust while maintaining the speed of 1.3 VSR. The longitudinal control force must not exceed 50 lbs throughout the maneuver without changing the trim control. 5.2.2.5 Longitudinal Control, Airspeed Variation—§ 25.145(b)(6). 5.2.2.5.1 Configuration.  Maximum landing weight or a lighter weight if considered more critical.  Most forward CG position.  Wing flaps extended to the maximum landing position.  Landing gear extended.  Engine power or thrust at flight idle. 5.2.2.5.2 Test Procedure. The airplane must be trimmed at a speed of 1.3 VSR. The speed should then be reduced to VSW and then increased to 1.6 VSR, or the maximum flap extended speed, VFE, whichever is lower. The longitudinal control force must not be greater than 50 lbs. Data from the static longitudinal stability tests in the landing configuration at forward CG, § 25.175(d), may be used to show compliance with this requirement. 5.2.2.6 Longitudinal Control, Flap Retraction and Power or Thrust Application—§ 25.145(c). 5.2.2.6.1 Configuration.  Critical combinations of maximum landing weights and altitudes.  Critical CG position.  Wing flaps extended to the maximum landing position and gated position, if applicable.  Landing gear extended.  Engine power or thrust for level flight at a speed of 1.08 VSR for propeller driven airplanes, or 1.13 VSR for turbojet powered airplanes. 5.2.2.6.2 Test procedure. With the airplane stable in level flight at a speed of 1.08 VSR for propeller driven airplanes, or 1.13 VSR for turbojet powered airplanes, retract the flaps to the full up position, or the next gated position, while simultaneously setting go-around power or thrust. Use the same power or thrust as is used to comply with the performance requirement of 5-19 ### Page 137 **AC 25-7D CHG 1** § 25.121(d), as limited by the applicable controllability requirements. It must be possible, without requiring exceptional piloting skill, to prevent losing altitude during the maneuver. Trimming is permissible at any time during the maneuver. If gates are provided, conduct this test beginning from the maximum landing flap position to the first gate, from gate to gate, and from the last gate to the fully retracted position. If there is a gate between any landing position and its associated go-around position(s), this test should also be conducted from that landing position through the gate to the associated go-around position. If there is more than one associated go-around position, this additional test should be conducted using the goaround position corresponding to the most retracted flap position, unless another position is more critical. Keep the landing gear extended throughout the test. 5.2.2.7 Longitudinal Control, Out-of-Trim Takeoff Conditions— §§ 25.107(e)(4) and 25.143(a)(1). See paragraphs 4.2.9.3.3 and 4.2.9.3.4 of this AC. 5.3 Directional and Lateral Control—§ 25.147. 5.3.1 Explanation. 5.3.1.1 Sections 25.147(a) and (b) provide criteria to determine if the airplane may have dangerous characteristics such as rudder lock or loss of directional control if it is maneuvered only with the rudder, while maintaining wings level, when one or two critical engines are inoperative. Some yaw capability into the operating engine(s) should be possible. It should also be possible to make reasonably sudden heading changes of up to 15°, as limited by rudder force or deflection, toward the inoperative engine(s). The intention of the requirement is that the airplane can be yawed as prescribed without needing to bank the airplane. Small variations of bank angle that are inevitable in a realistic flight test demonstration are acceptable. 5.3.1.2 Sections 25.147(c) and (e) require an airplane to be easily controllable with the critical engine(s) inoperative. Section 25.147(d) further requires that lateral control be sufficient to provide a roll rate necessary for safety, without excessive control forces or travel, at the speeds likely to be used with one engine inoperative. Compliance can normally be demonstrated in the takeoff configuration at V2 speed, because this condition is usually the most critical. Normal operation of a yaw stability augmentation system (SAS) should be considered in accordance with the normal operating procedures. Roll response with all engines operating, § 25.147(f), should be satisfactory for takeoff, approach, landing, and high-speed configurations. Any permissible configuration that could affect roll response should be evaluated. 5-20 ### Page 138 5.3.2 Procedures. The following test procedures outline an acceptable means for demonstrating compliance with § 25.147. 5.3.2.1 Directional Control: General—§ 25.147(a). 5.3.2.1.1 Configuration.  Maximum landing weight.  Most aft CG position.  Wing flaps extended to the approach position.  Landing gear retracted.  Yaw SAS on, and off if applicable.  Operating engine(s) at the power or thrust for level flight at 1.3 VSR, but not more than maximum continuous power or thrust.  Inoperative engine that would be most critical for controllability, with the propeller (for propeller airplanes) feathered. 5.3.2.1.2 Test Procedure. The airplane must be trimmed in level flight at the most critical altitude in accordance with § 25.21(c). Make heading changes into and away from the inoperative engine of up to 15° (not using more than 150 lbs rudder force), using the roll controls to maintain approximately wings level flight. The airplane should be controllable and free from any hazardous characteristics during this maneuver. For airplanes equipped with a rudder boost system, the evaluation should be done without rudder boost if the boost system can be inoperative. 5.3.2.2 Directional Control: Four or More Engines—§ 25.147(b). 5.3.2.2.1 Configuration.  Maximum landing weight.  Most forward CG position.  Wing flaps in the most favorable climb position (normally retracted).  Landing gear retracted.  Yaw SAS on, and off, as applicable.  Operating engines at the power or thrust required for level flight at 1.3 VSR1, but not more than maximum continuous power or thrust. 5-21 ### Page 139  Two inoperative engines that would be most critical for controllability with (if applicable) propellers feathered. 5.3.2.2.2 Test Procedure. The procedure outlined in paragraph 5.3.2.1.2 above is applicable to this test. 5.3.2.3 Lateral Control: General—§ 25.147(c). 5.3.2.3.1 Configuration.  Maximum takeoff weight.  Most aft CG position.  Wing flaps in the most favorable climb position.  Landing gear retracted and extended.  Yaw SAS on, and off, as applicable.  Operating engine(s) at maximum continuous power or thrust.  The inoperative engine that would be most critical for controllability, with the propeller (for propeller airplanes) feathered. 5.3.2.3.2 Test Procedure. With the airplane trimmed at 1.3 VSR1, turns with a bank angle of 20° must be demonstrated with and against the inoperative engine from a steady climb at 1.3 VSR1. It should not take exceptional piloting skill to make smooth, predictable turns. 5.3.2.4 Lateral Control: Roll Capability—§ 25.147(d). 5.3.2.4.1 Configuration.  Maximum takeoff weight.  Most aft CG position.  Wing flaps in the most critical takeoff position.  Landing gear retracted.  Yaw SAS on, and off, as applicable.  Operating engine(s) at maximum takeoff power or thrust.  The inoperative engine that would be most critical for controllability, with propellers (for propeller airplanes) feathered. 5.3.2.4.2 Test Procedure. With the airplane in trim, or as nearly as possible in trim, for straight flight at V2, establish a steady 30° banked turn. Demonstrate that the airplane 5-22 ### Page 140 can be rolled to a 30° bank angle in the other direction in not more than 11 seconds. The rudder may be used to the extent necessary to minimize sideslip. Demonstrate this maneuver in the most adverse direction. The maneuver may be unchecked, that is, the pilot need not apply a control input to stop the roll until after the 30° bank angle is achieved. Care should be taken to prevent excessive sideslip and bank angle during the recovery. 5.3.2.5 Lateral Control: Four or More Engines—§ 25.147(e). 5.3.2.5.1 Configuration.  Maximum takeoff weight.  Most aft CG position.  Wing flaps in the most favorable climb position.  Landing gear retracted and extended.  Yaw SAS on, and off, as applicable.  Operating engines at maximum continuous power or thrust.  Two inoperative engines most critical for controllability, with propellers (for propeller airplanes) feathered. 5.3.2.5.2 Test Procedure. The procedure outlined in paragraph 5.3.2.1.2 is applicable to this test. 5.3.2.6 Lateral Control: All Engines Operating—§ 25.147(f). 5.3.2.6.1 Configuration. All configurations within the flight envelope for normal operation. 5.3.2.6.2 Test Procedure. This is primarily a qualitative evaluation that should be conducted throughout the test program. Roll performance should be investigated throughout the flight envelope, including speeds to VFC/MFC, to ensure adequate peak roll rates for safety, considering the flight condition, without excessive control force or travel. Roll response during sideslips expected in service should provide maneuvering capabilities adequate to recover from such conditions. Approach and landing configurations should be carefully evaluated to ensure adequate control to compensate for gusts and wake turbulence while in close proximity to the ground. 5-23 ### Page 141 5.4 Minimum Control Speed—§ 25.149. 5.4.1 Explanation. 5.4.1.1 General. Section 25.149 defines requirements for minimum control speeds during takeoff climb (VMC), during takeoff ground roll (VMCG), and during approach and landing (VMCL and VMCL-2). The VMC (commonly referred to as VMCA) requirements are specified in § 25.149(a), (b), (c) and (d); the VMCG requirements are described in § 25.149(e); and the VMCL and VMCL-2 requirements are covered in § 25.149(f), (g) and (h). Section 25.149(a) states, “...the method used to simulate critical engine failure must represent the most critical mode of powerplant failure with respect to controllability expected in service.” That is, the power or thrust loss from the inoperative engine must be at the rate that would occur if an engine suddenly became inoperative in service. Prior to amendment 25-42 to § 25.149, the regulation required that rudder control forces must not exceed 180 lbs. With the adoption of amendment 25-42, rudder control forces became limited to 150 lbs. The relationships between VEF, V1, and VMCG are discussed in paragraph 4.2, Takeoff and Takeoff Speeds— §§ 25.105 and 25.107, and paragraph 4.3, Accelerate-Stop Distance— § 25.109. 5.4.1.2 Safety Concerns Addressed by VMCA. When flying with an inoperative engine, the asymmetric yawing moment must be compensated by aerodynamic forces created by rudder deflection and sideslip. When the speed decreases, sideslip increases rapidly in a non-linear manner. The purpose of the VMCA requirement is to ensure the airplane remains safely controllable with the maximum power or thrust asymmetry at any speed down to VMCA. 5.4.1.3 Weight Effect on VMCA. To maintain straight flight with an inoperative engine, as required by § 25.149(b), the lateral aerodynamic forces resulting from sideslip and rudder deflection must be balanced by the lateral component of weight (i.e., W * sin (bank angle)). The bank angle necessary to maintain straight flight is therefore approximately inversely proportional to the weight. Since § 25.149(b) allows VMCA to be determined with up to 5° of bank angle, this introduces a weight effect on VMCA. The heavier the weight, the lower the VMCA, but the greater will be the demonstrated sideslip. As an example, flying a heavy airplane at a VMCA speed determined at a lighter weight will result in the same sideslip, but a smaller bank angle (e.g., 4° instead of 5° if the airplane is 25 percent heavier). 5-24 ### Page 142 5.4.2 Procedures: General. 5.4.2.1 Prior to beginning the minimum control speed tests, the applicant should verify which engine’s failure will result in the largest asymmetric yawing moment (i.e., the “critical” engine). This is typically done by setting one outboard engine to maximum power or thrust, setting the corresponding opposite engine at idle, and decelerating with wings level until full rudder is required. By alternating power or thrust on/power or thrust off from left to right, the critical engine can be defined as the idle engine that requires the highest minimum speed to maintain a constant heading with full rudder deflection. 5.4.2.2 For propeller-driven airplanes, VMCA, VMCG, and VMCL (and VMCL-2, as applicable) should be determined by rendering the critical engine(s) inoperative and allowing the propeller to attain the position it automatically assumes. However, for some engine/propeller installations, a more critical drag condition could be produced as the result of a failure mode that results in a partial power condition that does not activate the automatic propeller drag reduction system (e.g., autofeather system). One example is a turbopropeller installation that can have a fuel control failure, which causes the engine to go to flight idle, resulting in a higher asymmetric yawing moment than would result from an inoperative engine. In such cases, in accordance with § 25.149(a), the minimum control speed tests must be conducted using the most critical failure mode. For propeller-driven airplanes where VMCA is based on operation of a propeller drag reduction system, VMCA should also be defined with the critical engine at idle to address the training situation where engine failure is simulated by retarding the critical engine to idle. If VMCA at idle is more than one knot greater than for the engine failure with an operating drag reduction system, the idle engine VMCA should be included in the normal procedures section of the AFM as advisory information to maintain the level of safety in the aforementioned training situation. 5.4.2.3 AFM values of VMCA, VMCG, and VMCL (and VMCL-2, as applicable) should be based on the maximum net power or thrust reasonably expected for a production engine. These speeds should not be based on specification power or thrust, since this value represents the minimum power or thrust guaranteed by the engine manufacturer, and the resulting minimum control speeds will not be representative of what could be achieved in operation. The maximum power or thrust used for scheduled AFM minimum control speeds should represent the high side of the tolerance band but may be determined by analysis instead of tests. 5.4.2.4 When determining VMCA, VMCL and VMCL-2, consideration should be given to the adverse effect of maximum approved lateral fuel imbalance on lateral control availability. This is especially of concern if tests or analysis 5-25 ### Page 143 show that the lateral control available is the determining factor of a particular VMC. 5.4.2.5 For changes to approved designs, the effect of any aerodynamic or propulsive changes on compliance with 25.149 must be assessed per § 21.20. For example, for design changes involving an increase in engine thrust, the effect of the higher thrust on minimum control speeds must be specifically evaluated, and, if found to be not negligible, must be accounted for. 5.4.3 Procedures: Minimum Control Speeds—Air (VMCA). 5.4.3.1 In showing compliance with the VMCA requirements, the following two conditions should be satisfied: 5.4.3.1.1 The stabilized (static) condition where constant heading is maintained without exceeding a 5° bank angle, and 5.4.3.1.2 The dynamic condition in which control is maintained without exceeding a heading change of 20°. Note: Separate tests are usually conducted to show compliance with these two conditions. 5.4.3.2 Static Test Procedure and Required Data. 5.4.3.2.1 To determine VMCA, use the configuration specified in § 25.149, except that VMCA is normally determined at minimum weight in order to minimize the stall speed and because static VMCA decreases with increased weight if a 5° bank angle is used. The requirement of § 25.149(c) that VMCA not exceed 1.13 VSR is based on VSR at maximum sea level takeoff weight. With the critical engine inoperative, the corresponding opposite engine should be adjusted to maximum takeoff power/thrust, and the airspeed decreased until heading can just be maintained with full rudder and no more than a 5° bank into the operating engine. For airplanes with more than two engines, the inboard engine(s) may be set to any power or thrust necessary to assist in developing the desired level of asymmetric power or thrust, or to achieve the desired flight path angle (normally level flight). 5.4.3.2.2 If the maximum asymmetric power or thrust that is permitted by the AFM operating limitations was maintained at the test day VMCA, and the rudder pedal force did not exceed the limit specified in § 25.149(d), the resulting speed may be used as the single value of VMCA for the airplane. If, at the option of the applicant, the AFM value of VMCA is to vary with pressure altitude and temperature, the test day minimum control speed and the corresponding power or thrust should be used to calculate an equivalent yawing moment coefficient (CN). This CN value may then be used to 5-26 ### Page 144 calculate VMCA as a function of takeoff power or thrust, thus permitting VMCA to be scheduled as a function of pressure altitude and temperature for takeoff data expansion and presentation in the AFM. (See appendix F of this AC for further discussion of VMCA correction.) 5.4.3.2.3 If maximum allowable takeoff power or thrust could not be developed at the flight test conditions, but maximum rudder deflection was achieved, then the VMCA value corresponding to sea level standard day maximum asymmetric power or thrust may be calculated from the CN attained at the test value of VMCA. Extrapolation using this constant CN method should be limited to 5 percent of the test day asymmetric power or thrust and should only be permitted if the rudder pedal force at the test day VMCA was not more than 95 percent of the limit value specified in § 25.149(d). For extrapolation beyond 5 percent power or thrust, a more rigorous analysis, using all the applicable stability and control terms, should be made. (See appendix F of this AC for further discussion of VMCA correction.) 5.4.3.2.4 If VMCA could not be achieved due to stall buffet, or excessive rudder pedal force, a parametric investigation should be undertaken to determine whether VMCA is limited by stall speed, maximum rudder deflection, or maximum allowable rudder pedal force. (See appendix G of this AC.) 5.4.3.3 Dynamic Test Procedures and Required Data. 5.4.3.3.1 After the static VMCA tests have been completed, dynamic engine cuts should be evaluated at a series of decreasing airspeeds to show that sudden engine failure at any speed down to the static VMCA value meets the requirements of § 25.149. The dynamic VMCA test is conducted by applying the maximum approved power/thrust to all outboard engines, stabilizing at the test airspeed, and then cutting fuel to the critical engine. The pilot must be able to recover to a straight flight condition (constant heading) with an angle of bank of not more than 5°.  Without deviating more than 20 degrees from the original heading,  While maintaining the test airspeed, without reducing power/thrust on the operating engine(s), and  Without exceeding the rudder pedal force limit of § 25.149(d). 5.4.3.3.2 In accordance with § 25.149(d), the airplane may not assume any dangerous attitude, nor require exceptional piloting skill, alertness, or strength. The maximum bank angle achieved during the tests may exceed 5° provided the airplane characteristics comply with this qualitative requirement. If the dynamic tests result in a VMCA greater than the static value, the increment between the static and dynamic VMCA at the same altitude should be added to the sea level extrapolated value. If the dynamic value is less than the static value, the static VMCA should be used for the AFM data expansion. 5-27 ### Page 145 5.4.3.3.3 If static VMCA is near stall speed at the minimum practicable test weight, or if the thrust-to-weight ratio (T/W) results in a trimmed pitch attitude of more than 20°, it is not feasible to attempt to accurately define a quantitative value of VMCA using a sudden engine cut because of the dynamics of the rapid pitch down maneuver required, and the hazard associated with a potential spin entry. Additionally, an extreme nose up attitude followed by an engine cut is not representative of an operational takeoff engine failure. Since § 25.107(e)(1)(ii) requires VR to be not less than 1.05 VMCA, and there is some additional speed increase prior to lift off, a transport airplane is typically never airborne below approximately 1.08 VMCA. Therefore, instead of using the dynamic method to define VMCA for these aircraft with high T/W or stall speed coincident with VMCA, it is more appropriate for a dynamic engine cut to be evaluated only for acceptable controllability, and at a more representative speed. For these airplanes, a dynamic engine cut should be evaluated at an airspeed of either 1.08 VSR or 1.1 VMCA (static), whichever is greater. During the entry to, and recovery from this maneuver, all the requirements of § 25.149(d) must be met. 5.4.3.3.4 For airplanes with rudder travel-limited VMCA that have increased power or thrust engines installed, with no changes to the airframe’s geometric layout or dimensions, it may not be necessary to conduct dynamic VMCA flight testing if the power or thrust has not increased more than 10 percent above the level at which dynamic VMCA had previously been demonstrated. (See appendix F of this AC.) 5.4.4 Procedures: Minimum Control Speed-Ground (VMCG)—§ 25.149(e). 5.4.4.1 It must be demonstrated that, when the critical engine is suddenly made inoperative at VMCG during the takeoff ground roll, the airplane is safely controllable if the takeoff is continued. During the demonstration, the airplane must not deviate more than 30 feet (25 feet prior to amendment 25-42) from the pre-engine-cut projected ground track. The critical engine) for ground minimum control speed testing should be determined during the takeoff ground run using techniques similar to these described in paragraph 5.4.2. If there is a significant difference in left and right rudder deflection, the loss of asymmetric propeller disc loading, due to near zero angle-of-attack during the takeoff roll, could result in the critical engine being on the opposite side of the airplane relative to the airborne minimum control speed tests. 5.4.4.2 Work up tests may be conducted by abruptly retarding the critical engine to idle to determine the airplane asymmetric control characteristics and provide data from which an estimate of VMCG can be made. Due to the engine spindown characteristics with the critical engine retarded to idle, the speed will not, in general, be representative of the VMCG speed that would be obtained with a fuel cut. Therefore, the certification tests for 5-28 ### Page 146 VMCG should be conducted using fuel cuts. Starting from a speed comfortably above the estimated VMCG and with the maximum takeoff power or thrust level to be certified, several fuel cuts should be made at decreasing calibrated airspeeds to establish the minimum airspeed at which the lateral deviation is less than or equal to 30 feet. VMCG is determined for zero crosswind conditions. However, in light crosswind test conditions the VMCG value determined should be that which is appropriate to the adverse crosswind or, at the applicant’s option, may be corrected to a zero crosswind value using runs made on reciprocal headings. 5.4.4.3 During determination of VMCG, engine failure recognition should be provided by the pilot:  Feeling a distinct change in the directional tracking characteristics of the airplane, or  Seeing a directional divergence of the airplane with respect to the view outside the airplane. 5.4.4.4 Directional control of the airplane should be accomplished by use of the rudder only. All other controls, such as ailerons and spoilers, should only be used to correct any alterations in the airplane attitude and to maintain a wings level condition. Pilot input to controls to supplement the rudder effectiveness should not be used. Care should also be taken not to inadvertently apply brake pressure during large rudder deflections, as this will invalidate the test data. 5.4.4.5 VMCG testing should be conducted at the most critical weight in the range where VMCG may impact AFM V1 speeds. 5.4.4.6 VMCG testing should be conducted at aft CG and with the nose wheel free to caster, to minimize the stabilizing effect of the nose gear. If the nose wheel does not caster freely, the test may be conducted with enough nose up elevator applied to lift the nose wheel off the runway. 5.4.4.7 VMCG testing should not be conducted on runways with excessive crowning (i.e., cross-runway slope) unless the effects of such crowning are determined to be conservative. 5.4.4.8 For airplanes with certification bases prior to amendment 25-42, VMCG values may be demonstrated with nose wheel rudder pedal steering operative for dispatch on wet runways. The test should be conducted on an actual wet, smooth (i.e., not grooved or PFC) runway. The test(s) should include engine failure at or near a minimum VEF associated with minimum VR to demonstrate adequate controllability during rotation, liftoff, and the initial climbout. The VMCG values obtained by this method are applicable for wet or dry runways only, not for icy runways. 5-29 ### Page 147 5.4.5 Procedures: Minimum Control Speed during Approach and Landing (VMCL)— § 25.149(f). 5.4.5.1 This section is intended to ensure that the airplane is safely controllable following an engine failure during an all-engines-operating approach and landing. From a controllability standpoint, the most critical case usually consists of an engine failing after the power or thrust has been increased to perform a go-around from an all-engines-operating approach. Section 25.149(f) requires the minimum control speed to be determined that allows a pilot of average skill and strength to retain control of the airplane after the critical engine becomes inoperative and to maintain straight flight with less than 5° of bank angle. Section 25.149(h) requires that sufficient lateral control be available at VMCL to roll the airplane through an angle of 20°, in the direction necessary to initiate a turn away from the inoperative engine, in not more than five seconds when starting from a steady straight flight condition. 5.4.5.2 Conduct this test using the most critical of the all-engines-operating approach and landing configurations or, at the option of the applicant, each of the all-engines-operating approach and landing configurations. The procedures given in paragraphs 5.4.3.2 and 5.4.3.2.3 for VMCA may be used to determine VMCL, except that flap and trim settings should be appropriate to the approach and landing configurations, the power or thrust on the operating engine(s) should be set to the go-around power or thrust setting, and compliance with all VMCL requirements of § 25.149(f) and (h) must be demonstrated. 5.4.5.3 In accordance with § 25.149(f)(5) for propeller driven airplanes, the propeller must be in the position it achieves without pilot action following engine failure, assuming the engine fails while at the power or thrust necessary to maintain a 3° approach path angle. 5.4.5.4 At the option of the applicant, a one-engine-inoperative landing minimum control speed, VMCL(1 out), may be determined in the conditions appropriate to an approach and landing with one engine having failed before the start of the approach. In this case, only those configurations recommended for use during an approach and landing with one engine inoperative need be considered. The propeller of the inoperative engine, if applicable, may be feathered throughout. The resulting value of VMCL(1 out) may be used in determining the recommended procedures and speeds for a one-engine-inoperative approach and landing. 5.4.6 Procedures: Minimum Control Speed with Two Inoperative Engines during Approach and Landing (VMCL-2)—§ 25.149(g). 5.4.6.1 For airplanes with three or more engines, VMCL-2 is the minimum speed for maintaining safe control during the power or thrust changes that are likely 5-30 ### Page 148 to be made following the failure of a second critical engine during an approach initiated with one engine inoperative. 5.4.6.2 In accordance with § 25.149(g)(5) for propeller driven airplanes, the propeller of the engine that is inoperative at the beginning of the approach may be in the feathered position. The propeller of the more critical engine must be in the position it automatically assumes following engine failure. 5.4.6.3 Conduct this test using the most critical approved one-engine-inoperative approach or landing configuration (usually the minimum flap deflection), or at the option of the applicant, each of the approved one-engine-inoperative approach and landing configurations. The following demonstrations should be conducted to determine VMCL-2: 5.4.6.3.1 With the power or thrust on the operating engines set to maintain a -3° glideslope with one critical engine inoperative, the second critical engine is made inoperative, and the remaining operating engine(s) are advanced to the go-around power or thrust setting. The VMCL-2 speed is established by the procedures presented in paragraphs 5.4.3.2 and 5.4.3.3 below for VMCA, except that flap and trim settings should be appropriate to the approach and landing configurations, the power or thrust on the operating engine(s) should be set to the go-around power or thrust setting, and compliance with all VMCL-2 requirements of § 25.149(g) and (h) must be demonstrated. 5.4.6.3.2 With power or thrust on the operating engines set to maintain a -3° glideslope, with one critical engine inoperative: 1. Set the airspeed at the value determined in paragraph 5.4.6.3.1 above and, with zero bank angle, maintain a constant heading using trim to reduce the control force to zero. If full trim is insufficient to reduce the control force to zero, full trim should be used plus control deflection as required; and 2. Make the second critical engine inoperative and retard the remaining operating engine(s) to minimum available power or thrust without changing the directional trim. The VMCL-2 determined in paragraph 5.4.6.3.1 is acceptable if constant heading can be maintained without exceeding a 5° bank angle and the limiting conditions of § 25.149(h). 3. Starting from a steady straight flight condition, demonstrate that sufficient lateral control is available at VMCL-2 to roll the airplane through an angle of 20° in the direction necessary to initiate a turn away from the inoperative engines in not more than five seconds. This maneuver may be flown in a bank-to-bank roll through a wings level attitude. 5-31 ### Page 149 5.4.6.4 At the option of the applicant, a two-engines-inoperative landing minimum control speed, VMCL-2(2 out), may be determined in the conditions appropriate to an approach and landing with two engines having failed before the start of the approach. In this case, only those configurations recommended for use during an approach and landing with two engines inoperative need be considered. The propellers of the inoperative engines, if applicable, may be feathered throughout. The values of VMCL-2 or VMCL-2(2 out) should be used as guidance in determining the recommended procedures and speeds for a two-engines-inoperative approach and landing. 5.4.6.5 Autofeather Effects. Where an autofeather or other drag limiting system is installed, and will be operative at approach power settings, its operation may be assumed in determining the propeller position achieved when the engine fails. Where automatic feathering is not available, the effects of subsequent movements of the engine and propeller controls should be considered, including fully closing the power lever of the failed engine in conjunction with maintaining the go-around power setting on the operating engine(s). 5-32 ### Page 150 CHAPTER 6. FLIGHT: TRIM 6.1 Trim—§ 25.161. 6.1.1 Explanation. Adequate trim capability should be provided for any flight condition that it is reasonable to assume will be maintained steadily for any appreciable time. 6.1.2 Procedures. 6.1.2.1 The trim requirements specify the ranges of speed and the airplane configurations at which the airplane must be able to maintain trim. 6.1.2.2 All weights, from the minimum in-flight weight to the maximum takeoff weight, should be considered. For airplanes with unpowered controls, the lowest weight may be more critical since this results in the lowest airspeed. 6.2 [Reserved.] 6-1 ### Page 151 CHAPTER 7. FLIGHT: STABILITY 7.1 General—§ 25.171. [Reserved] 7.2 Static Longitudinal Stability and Demonstration of Static Longitudinal Stability— §§ 25.173 and 25.175. 7.2.1 Explanation. 7.2.1.1 Static Longitudinal Stability—§ 25.173. 7.2.1.1.1 Compliance with the general requirements of § 25.173 is determined from a demonstration of static longitudinal stability under the conditions specified in § 25.175. 7.2.1.1.2 The requirement is to have a pull force to obtain and maintain speeds lower than trim speed, and a push force to obtain and maintain speeds higher than trim speed. There may be no force reversal at any speed that can be obtained, except lower than the minimum for steady, unstalled flight or, higher than the landing gear or wing flap operating limit speed or VFC/MFC, whichever is appropriate for the test configuration. The required trim speeds are specified in § 25.175. 7.2.1.1.3 When the control force is slowly released from any speed within the required test speed range, the airspeed must return to within 10 percent of the original trim speed in the climb, approach, and landing conditions, and return to within 7.5 percent of the trim speed in the cruising condition specified in § 25.175 (free return). 7.2.1.1.4 The average gradient of the stick force versus speed curves for each test configuration may not be less than 1 pound for each 6 knots for the appropriate speed ranges specified in § 25.175. Therefore, after each curve is drawn, draw a straight line from the intersection of the curve and the required maximum speed to the trim point. Then draw a straight line from the intersection of the curve and the required minimum speed to the trim point. The slope of these lines must be at least 1 pound for each 6 knots. The local slope of the curve must remain stable for this range. 7.2.1.2 Demonstration of Static Longitudinal Stability—§ 25.175. This section specifically defines the flight conditions, airplane configurations, trim speed, test speed ranges, and power or thrust settings to be used in demonstrating compliance with the longitudinal stability requirements. 7-1 ### Page 152 7.2.2 Procedures. 7.2.2.1 Stabilized Method. 7.2.2.1.1 For the demonstration of static longitudinal stability, the airplane should be trimmed in smooth air at the conditions required by the regulation. Aft CG loadings are generally most critical. After stabilizing at the trim speed, apply a light pull force and stabilize at a slower speed. Continue this process in increments, the size of the speed increment being dependent on the speed spread being investigated, until reaching the minimum speed for steady, unstalled flight or the minimum speed appropriate for the configuration. A continuous pull force should be used from the trim speed on each series of test points to eliminate hysteresis effects. At the end of the required speed range, the force should be gradually relaxed to allow the airplane to return slowly toward the trim speed and zero stick force. Depending on the amount of friction in the control system, the eventual speed at which the airplane stabilizes will normally be less than the original trim speed. The new speed, called the free return speed, must meet the requirements of § 25.173. 7.2.2.1.2 Starting again at the trim speed, and with the airplane in trim, push forces should be gradually applied and gradually relaxed in the same manner as described in paragraph 7.2.2.1.1 above. 7.2.2.1.3 The above techniques result in several problems in practice. One effect of changing airspeed is a change of altitude, with a corresponding change in Mach number and power or thrust output. Consequently, a reasonably small altitude band, limited to ±3,000 feet, should be used for the complete maneuver. If this altitude band is exceeded, regain the original trim altitude by changing the power or thrust setting and flap and gear position as necessary, but without changing the trim setting. Then continue the push or pull maneuver in the original configuration. Testing somewhat beyond the required speed limits in each direction assures that the resulting data covers at least the required speed ranges. It will also be noted in testing that while holding force constant at each data point, the airspeed and instantaneous vertical speed vary in a cyclic manner. This is due to the long period (phugoid) oscillation. Care should be exercised in defining and evaluating the data point, since it may be biased by this phugoid oscillation. Averaging these oscillating speeds at each data point is an acceptable method of eliminating this effect. Extremely smooth air improves the quality of the test data. In-bay and cross-bay wing fuel shift is another issue experienced in some airplanes. In-bay fuel shift occurs rapidly with pitch angle; therefore, consideration should be given to testing with fuel loadings that provide the maximum shift since it is generally destabilizing. Slower, cross-bay fuel shift, or burn from an aft tank, can influence the measured stability but usually only because of the time required to obtain the data points. This testing induced instability 7-2 ### Page 153 05/04/18 AC 25-7D should be removed from the data before evaluating the slope of the stick force versus speed. 7.2.2.2 Acceleration-Deceleration Method. 7.2.2.2.1 Trim at the desired airspeed and note the power or thrust setting. Without changing pitch trim, increase power or thrust to accelerate the airplane to the extreme speed of the desired data band. Using elevator control as needed, maintain approximately a constant altitude. Then, without changing pitch trim, quickly reset the power or thrust to the original power setting and allow the airplane to decelerate at a constant altitude back to the original trim speed. Obtain longitudinal static stability data during the deceleration to trim speed with the power and the pitch trim position the same as the original trim data point. 7.2.2.2.2 Obtain data below the trim speed in a similar manner, by reducing power or thrust to decelerate the airplane to the lowest speed in the data band. Using elevator control as needed without changing pitch trim, maintain approximately a constant altitude. Then, without changing pitch trim, quickly reset the power to the original power setting, and record the data during the level flight acceleration back to trim speed. If, because of thrust/drag relationships, the airplane has difficulty returning to the trim conditions, small altitude changes within ±2,000 feet can also be used to coax an airplane back to trim speed. Level flight is preferred, if possible. Obtain speed and elevator stick force data approximately every 10 knots of speed change. 7.2.2.2.3 The resulting pilot longitudinal force test points should be plotted versus airspeed to show the positive stable gradient of static longitudinal stability and that there are no “local” reversals in the stick force versus airspeed relationship over the range of airspeeds tested. This plot should also show the initial trim point and the two return-to-trim points to evaluate the return-to-trim characteristics. (See figure 7-1 below.) 7-3 ### Page 154 Figure 7-1. Longitudinal Static Stability 7.2.2.3 Examples of “local reversals” are given in figure 7-2 below. Curves A and C depict a local gradient reversal within the required speed range. Even though it might be argued that the “average gradient” meets the 1 pound in 6 knots criterion, the gradient reversals would render these characteristics unacceptable. Curve B depicts a situation in which the gradient reverses, but only outside the required speed range. In addition, Curve B demonstrates a situation in which the local gradient does not always meet the required 1 pound in 6 knots, even though the average gradient does. 7-4 ### Page 155 05/04/18 AC 25-7D Figure 7-2. Local Reversal 7.3 Static Directional and Lateral Stability—§ 25.177. 7.3.1 Explanation. 7.3.1.1 Static Directional Stability. Positive static directional stability is defined as the tendency to recover from a skid with the rudder free. Positive static directional stability is required by § 25.177(a) for any landing gear and flap position and symmetrical power or thrust condition at speeds from 1.13 VSR1 up to VFE, VLE, or VFC/MFC, as appropriate for the airplane configuration. 7.3.1.2 Static Lateral Stability. Positive static lateral stability is defined as the tendency to raise the low wing in a sideslip with hands off the roll controls. Section 25.177(b) requires that static lateral stability not be negative in any landing gear and flap position and symmetrical power or thrust condition at speeds from 1.13 VSR1 to VFE, VLE, or VMO/MMO, as appropriate for the airplane configuration. At speeds from VMO/MMO to VFC/MFC, negative static lateral stability is permitted by § 25.177(b), if the divergence is:  Gradual,  Easily recognizable by the pilot, and 7-5 ### Page 156  Easily controllable by the pilot. 7.3.1.3 Steady Straight Sideslips. 7.3.1.3.1 Section 25.177(c) requires, in steady, straight sideslips throughout the range of sideslip angles appropriate to the operation of the airplane, that the aileron and rudder control movements and forces be proportional to the angle of sideslip. Also, the factor of proportionality must lie between limits found necessary for safe operation. Section 25.177(c) also states th at that the range of sideslip angles evaluated mu st include those sideslip angles resulting from the lesser of: (1) one-half of the available rudder control input; and (2) a rudder control force of 180 lbs. This means that if using one-half of the available rudder control input takes less than 180 lbs of force, then compliance mu st be based on using one-half of the available rudder control input. If application of 180 lbs of rudder control force results in using less than one-half of the available rudder control input, then compliance must be based on applying 180 lbs of rudder control force. By cross-reference to § 25.177(a), § 25.177(c) requires that these steady, straight sideslip criteria be met for all landing gear and flap positions and symme tric al power or thrust conditions at speeds from 1.13 VSR1 to VFE, VLE, or VFC/MFC, as appropriate for the configuration. 7.3.1.3.2 Experience has shown that an acceptable metho d for determining th e appropriate sideslip angle for the operation of a transport category airplane is provided by the following equation: Where: 7.3.1.3.3 Recognizing that smaller sideslip angles are appropriate as speed is increased, this equation provides sideslip angle as a function of airspeed. The equation is based on the theoretical sideslip value for a 30-knot crosswind, but has been shown to conservatively represent (i.e., exceed) the sideslip angles achieved in maximum crosswind takeoffs and landings and minimum static and dynamic control speed testing for a variety of transport category airplanes. Experience has also shown that a maximum sideslip angle of 15° is generally appropriate for most transport category airplanes even though the equation above may provide a higher sideslip angle. However, limiting the maximum sideslip angle to 15° may not be appropriate for airplanes with low approach speeds or high crosswind capability. 7.3.1.3.4 A lower sideslip angle than that provided in paragraph 7.3.1.3.2 above may be used if it is substantiated that the lower value conservatively 7-6 ### Page 157 05/04/18 AC 25-7D covers all crosswind conditions, engine failure scenarios, and other conditions where sideslip may be experienced within the approved operating envelope. Conversely, a higher value should be used for airplanes where test evidence indicates that a higher value would be appropriate to the operation of the airplane. 7.3.1.3.5 For the purpose of showing compliance with the requirement out to sideslip angles associated with the lesser of (1) one half of the available rudder control input; and (2) a rudder control force of 180 lbs; there is no need to consider a rudder control input beyond that corresponding to full available rudder surface travel. Some rudder control system designs may limit the available rudder surface deflection such that full deflection for the particular flight condition, or the maximum commanded sideslip angle for the flight condition, is reached before the rudder control reaches onehalf of its available travel. In such cases, further rudder control input is unnecessary as it would not result in a higher sideslip angle, and therefore would not affect compliance with the rule. 7.3.1.4 Full Rudder Sideslips. 7.3.1.4.1 At sideslip angles greater than those appropriate for normal operation of the airplane, up to the sideslip angle at which full rudder control input is used or a rudder control force of 180 lbs is obtained, § 25.177(d) requires that the rudder pedal control may not reverse, and increased rudder deflection must be needed for increased angles of sideslip. The goals of this higher-than-normal sideslip angle test are to show that at full rudder control input, or at maximum expected pilot effort (1) the rudder control force does not reverse, and (2) increased rudder deflection must be needed for increased angles of sideslip, thus demonstrating freedom from rudder lock or fin stall, and adequate directional stability for maneuvers involving large rudder inputs. 7.3.1.4.2 Compliance with this requirement should be shown using straight, steady sideslips. However, if full lateral control input is reached before full rudder control travel or a rudder control force of 180 lbs is reached, the maneuver may be continued in a non-steady heading (i.e., rolling and yawing) maneuver. Care should be taken to prevent excessive bank angles that may occur during this maneuver. 7.3.1.4.3 Section 25.177(d) states that the criteria listed in paragraph 7.3.1.4.1 above must be met at all approved landing gear and flap positions for the range of operating speeds and power conditions appropriate to each landing gear and flap position with all engines operating. The range of operating speeds and power conditions appropriate to each landing gear and flap position with all engines operating should be consistent with the following: 7-7 ### Page 158 05/04/18 AC 25-7D 1. For takeoff configurations, speeds from V2+xx (airspeed approved for all-engines-operating initial climb) to VFE or VLE, as appropriate, and takeoff power/thrust; 2. For flaps up configurations, speeds from 1.23 VSR to VLE or VMO/MMO, as appropriate, and power from idle to maximum continuous power/thrust; 3. For approach configurations, speeds from 1.23 VSR to VFE or VLE, as appropriate, and power from idle to go-around power/thrust; and 4. For landing configurations, speeds from VREF-5 knots to VFE or VLE, as appropriate, with power from idle to go-around power/thrust at speeds from VREF to VFE/VLE, and idle power at VREF-5 knots (to cover the landing flare). 7.3.2 Procedures. The test conditions should include each flap and landing gear configuration as described in paragraphs 7.3.1.1 through 7.3.1.4 of this AC at an altitude appropriate to each configuration. 7.3.2.1 Basic Tests for Static Directional and Lateral Stability. 7.3.2.1.1 Static Directional Stability. To check static directional stability with the airplane in the desired configuration and stabilized at the trim speed, the airplane is slowly yawed in both directions while maintaining the wings level with the roll controls. When the rudder is released, the airplane should tend to return to straight flight. 7.3.2.1.2 Static Lateral Stability. To check lateral stability with a particular configuration and trim speed, conduct steady, straight sideslips at the trim speed by maintaining the airplane heading with rudder and banking with the roll controls. When the roll controls are released, with the rudder held fixed, the low wing should tend to return to level. Initial bank angle should be appropriate to type; however, it is recommended that it should not be less than 10° or that necessary to maintain the steady, straight sideslip with one-half rudder deflection, whichever occurs first. Roll control centering by the pilot should not be permitted during this evaluation. The intent of this testing is to evaluate the short-term response of the airplane; therefore, long-term effects, such as those due to spanwise fuel movement, need not be taken into account. 7.3.2.1.3 Steady, Straight Sideslips. Steady, straight sideslips should be conducted in each direction to show that the aileron and rudder control movements and forces are substantially proportional to the angle of sideslip in a stable sense, and that the factor of 7-8 ### Page 159 proportionality is within the limits found necessary for safe operation. These tests should be conducted at progressively greater sideslip angles up to the sideslip angle appropriate to the operation of the airplane (see paragraph 7.3.1.3.2) or the sideslip angle associated with one-half of the available rudder control input (as limited by a rudder control force of 180 lbs), whichever is greater. 7.3.2.1.4 When determining the rudder and aileron control forces, the controls should be relaxed at each point to find the minimum force needed to maintain the control surface deflection. If excessive friction is present, the resulting low forces will indicate the airplane does not have acceptable stability characteristics. 7.3.2.1.5 Instead of conducting each of the separate qualitative tests described in paragraph 7.3.2.1, the applicant may use recorded quantitative data showing aileron and rudder control force and position versus sideslip (left and right) to the appropriate limits in the steady heading sideslips conducted to show compliance with § 25.177(c). If the control force and position versus sideslip indicates positive dihedral effect and positive directional stability, compliance with § 25.177(a) and (b) will have been successfully demonstrated. 7.3.2.2 Full Rudder Sideslips. 7.3.2.2.1 Rudder lock is that condition where the rudder over-balances aerodynamically and either deflects fully with no additional pilot input or does not tend to return to neutral when the pilot input is released. It is indicated by a reversal in the rudder control force as sideslip angle is increased. Full rudder sideslips are conducted to determine the rudder control forces and deflections out to sideslip angles associated with full rudder control input (or as limited by a rudder control force of 180 lbs) to investigate the potential for rudder lock and lack of directional stability. 7.3.2.2.2 To check for positive directional stability and for the absence of rudder lock, conduct steady heading sideslips at increasing sideslip angles until obtaining full rudder control input or a rudder control force of 180 lbs. If full lateral control is reached before reaching the rudder control limit or 180 lbs of rudder control force, continue the test to the rudder limiting condition in a non-steady heading sideslip maneuver. 7.3.2.3 Control Limits. The control limits approved for the airplane should not be exceeded when conducting the flight tests required by § 25.177. 7.3.2.4 Flight Test Safety Concerns. In planning for and conducting the full rudder sideslips, items relevant to flight test safety should be considered, including: 7-9 ### Page 160 05/04/18 AC 25-7D  Inadvertent stalls,  Effects of sideslip on stall protection systems,  Actuation of stick pusher, including the effects of sideslip on angle-of-attack sensor vanes,  Heavy buffet,  Exceeding flap loads or other structural limits,  Extreme bank angles,  Propulsion system behavior (e.g., propeller stress, fuel and oil supply, and inlet stability),  Minimum altitude for recovery,  Resulting roll rates when the aileron limit is exceeded,  Position errors and effects on electronic or augmented flight control systems, especially when using the airplane’s production airspeed system, and  Rudder loads, particularly those that may occur with dynamic rudder inputs. 7.4 Dynamic Stability—§ 25.181. 7.4.1 Explanation. The dynamic stability tests described in this section should be conducted over the speed range of 1.13 VSR to VFE, VLE or VFC/MFC, as appropriate. 7.4.1.1 Dynamic Longitudinal Stability. 7.4.1.1.1 The short period oscillation is the first oscillation the pilot sees after disturbing the airplane from its trim condition with the pitch control (as opposed to the long period (phugoid)). Care should be taken that the control movement used to excite the motion is not too abrupt. 7.4.1.1.2 Heavily damped means that the oscillation has decreased to 1/10 the initial amplitude within approximately two cycles after completion of the control input. 7.4.1.1.3 Short period oscillations must be heavily damped, both with controls free and controls fixed. 7.4.1.2 Dynamic Lateral-Directional Stability. The evaluation of the dynamic lateral-directional stability should include any combined lateral-directional oscillation (“Dutch roll”) occurring over the speed range appropriate to the airplane configuration. This oscillation 7-10 ### Page 161 must be positively damped with controls free and must be controllable with normal use of the primary controls without requiring exceptional piloting skill. 7.4.2 Procedures. 7.4.2.1 Dynamic Longitudinal Stability. 7.4.2.1.1 The test for longitudinal dynamic stability is accomplished by a rapid movement or pulse of the longitudinal control in a nose up and nose down direction at a rate and degree necessary to obtain a short period pitch response from the airplane. 7.4.2.1.2 Dynamic longitudinal stability should be checked at a sufficient number of points in each configuration to assure compliance at all operational speeds. 7.4.2.2 Dynamic Lateral-Directional Stability. 7.4.2.2.1 A typical test for lateral-directional dynamic stability is accomplished by a rudder doublet input at a rate and amplitude that will excite the lateral-directional response (i.e., Dutch roll). The control input should be in phase with the airplane’s oscillatory response. 7.4.2.2.2 Dynamic lateral-directional stability should be checked under all conditions and configurations. If critical, special emphasis should be placed on adverse wing fuel loading conditions. 7.4.2.3 Airplanes Equipped with Stability Augmentation Systems (SAS). 7.4.2.3.1 In the event a SAS is required for the airplane to show compliance with § 25.181(a) or (b), it must meet the requirements of §§ 25.671 and 25.672. Additionally: 7.4.2.3.2 If the airplane is equipped with only one SAS (i.e., a single strand system), in accordance with § 25.672, compliance with the dynamic stability requirements of § 25.181(a) or (b), as applicable, must be shown throughout the normal operating flight envelope to be certificated with the SAS operating, and in a reduced, practical operating flight envelope that will permit continued safe flight and landing with the SAS inoperative. 7.4.2.3.3 If the airplane is equipped with more than one SAS, the resulting effects of SAS failure should be considered when determining whether or not the primary and any redundant SAS should be operating simultaneously for showing compliance with the dynamic stability requirements of § 25.181(a) or (b). If the primary and redundant SAS are dissimilar, the functional capability (i.e., control authority) of the redundant SAS should be considered with regard to restricting the operating envelope after failure of the primary SAS. At the applicant’s option, however, compliance with 7-11 ### Page 162 05/04/18 AC 25-7D § 25.181(a) or (b) may still be demonstrated to a reduced flight envelope with no SAS operating as described in paragraph 7.4.2.3.2 above. 7.4.2.3.4 Regardless of the SAS redundancy, the airplane should be safely controllable at the point of system failure or malfunction anywhere in the approved operating flight envelope of the airplane. Accordingly, it should be demonstrated that the airplane remains controllable during transition from the operating SAS to any redundant SAS, and during transition from anywhere in the normal operating envelope to the reduced practical operating envelope of § 25.672(c), if applicable. Airplane controllability should be demonstrated to meet the following levels as defined by the FAA HQRM. (The FAA HQRM is described in appendix E of this AC.) 1. In the normal operating flight envelope with the SAS operating, the handling qualities should be “satisfactory” (SAT) as defined by the FAA HQRM. 2. At the point of SAS failure in the normal operating envelope, the airplane should be “controllable” (CON), as defined by the FAA HQRM, during the short term transitory period required to attain a speed and configuration that will permit compliance with paragraph 3 below. 3. During transition from the primary SAS to a redundant SAS, or from the normal operating envelope to a reduced, practical operating envelope (where applicable), the handling qualities should be “adequate” (ADQ) as defined by the HQRM. 4. In the reduced, practical operating flight envelope that will permit continued safe flight and landing, the handling qualities should be “satisfactory” (SAT) as defined by the HQRM. 7-12 ### Page 163 05/04/18 AC 25-7D CHAPTER 8. FLIGHT: STALLS 8.1 Stall Testing. 8.1.1 Applicable 14 CFR Regulations.  Section 25.21(c), Proof of compliance.  Section 25.103, Stall speed.  Section 25.143, Controllability and maneuverability—General.  Section 25.201, Stall demonstration.  Section 25.203, Stall characteristics.  Section 25.207, Stall warning. 8.1.2 Explanation. 8.1.2.1 The purpose of stall testing is threefold: 8.1.2.1.1 To define the reference stall speeds and how they vary with weight, altitude, and airplane configuration. 8.1.2.1.2 To demonstrate that handling qualities are adequate to allow a safe recovery from the highest angle-of-attack attainable in normal flight (stall characteristics). 8.1.2.1.3 To determine that there is adequate pre-stall warning (either aerodynamic or artificial) to allow the pilot time to recover from any probable high angle-of-attack condition without inadvertently stalling the airplane. 8.1.2.2 During this testing, the angle-of-attack should be increased at least to the point where the behavior of the airplane gives the pilot a clear and distinctive indication through the inherent flight characteristics or the characteristics resulting from the operation of a stall identification device (e.g., a stick pusher) that the airplane is stalled. 8.1.3 Stall Demonstration—§ 25.201. 8.1.3.1 The airplane is considered to be fully stalled when any one or a combination of the characteristics listed below occurs to give the pilot a clear and distinctive indication to cease any further increase in angle-of-attack, at which time recovery should be initiated using normal techniques. 8.1.3.1.1 The pitch control reaches the aft stop and is held full aft for two seconds, or until the pitch attitude stops increasing, whichever occurs later. In the case of turning flight stalls, recovery may be initiated once the pitch 8-1 ### Page 164 05/04/18 AC 25-7D control reaches the aft stop when accompanied by a rolling motion that is not immediately controllable (provided the rolling motion complies with § 25.203(c)). 8.1.3.1.2 An uncommanded, distinctive, and easily recognizable nose down pitch that cannot be readily arrested. This nose down pitch may be accompanied by a rolling motion that is not immediately controllable, provided that the rolling motion complies with § 25.203(b) or (c), as appropriate. 8.1.3.1.3 The airplane demonstrates an unmistakable, inherent aerodynamic warning of a magnitude and severity that is a strong and effective deterrent to further speed reduction. This deterrent level of aerodynamic warning (i.e., buffet) should be of a much greater magnitude than the initial buffet ordinarily associated with stall warning. An example is a large transport airplane that exhibits “deterrent buffet” with flaps up and is characterized by an intensity that inhibits reading cockpit instruments and would require a strong determined effort by the pilot to increase the angle-of-attack any further. 8.1.3.1.4 The activation point of a stall identification device that provides one of the characteristics listed above. See paragraph 42.1 of this AC for additional guidance material on demonstrating compliance with the regulatory requirements of part 25 for stall identification systems. 8.1.3.2 It should be recognized that the point at which the airplane is considered stalled may vary, depending on the airplane configuration (e.g., flaps, gear, CG, and gross weight). In any case, the angle-of-attack should be increased until one or more of these characteristics is reached for all likely combinations of variables. 8.1.4 Stall Speeds. 8.1.4.1 Background. Since many of the regulations pertaining to performance and handling qualities specify trim speeds and other variables that are functions of stall speeds, it is desirable to accomplish the stall speed testing early in the program, so the data are available for subsequent testing. Because of this interrelationship between the stall speeds and other critical performance parameters, it is essential that accurate measurement methods be used. Most standard airplane pitot-static systems are unacceptable for stall speed determination. These tests require the use of properly calibrated instruments and usually require a separate test airspeed system. 8-2 ### Page 165 8.1.4.2 Configuration. 8.1.4.2.1 Stall speeds should be determined for all aerodynamic configurations to be certificated for use in the takeoff, en route, approach, and landing configurations. 8.1.4.2.2 The CG positions to be used should be those that result in the highest stall speeds for each weight (forward CG in most cases). 8.1.4.2.3 Sufficient testing should be conducted to determine the effects of weight on stall speed. Altitude effects (compressibility, Reynolds Number) may also be considered if credit for variations in these parameters is sought by the applicant. If stall speeds are not to be defined as a function of altitude, then all stall speed testing should be conducted at a nominal altitude no lower than 1,500 feet above the maximum approved takeoff and landing altitude. (See paragraph 8.1.4.5.7 of this AC.) 8.1.4.3 Procedures. 8.1.4.3.1 The airplane should be trimmed for hands-off flight at a speed 13 percent to 30 percent above the anticipated VSR, with the engines at idle and the airplane in the configuration for which the stall speed is being determined. Then, using only the primary longitudinal control for speed reduction, maintain a constant deceleration (entry rate) until the airplane is stalled, as defined in § 25.201(d) and paragraph 8.1.3.1 of this AC. Following the stall, engine power or thrust may be used as desired to expedite recovery. 8.1.4.3.2 A sufficient number of stalls (normally four to eight) should be accomplished at each critical combination of weight, altitude, CG, and external configuration. The intent is to obtain enough data to determine the stall speed at an entry rate not exceeding 1.0 knot/second. During the maneuver for determining stall speeds, the flight controls should be operated smoothly in order to achieve good data quality rather than trying to maintain a constant entry rate because experience has shown that adjusting the flight controls to maintain a constant entry rate leads to fluctuations in load factor and significant data scatter. 8.1.4.3.3 During the stall speed testing, the stall characteristics of the airplane must also satisfy the requirements of § 25.203(a) and (b). For airplanes that have stall identification devices for which the angle-of-attack for activation is biased by angle-of-attack rate, some additional considerations are necessary. The stall speeds are normalized against an average airspeed deceleration rate, as described in paragraph 8.1.4.5.5. However, stall identification systems generally activate at a specific angle-of-attack, biased by an instantaneous angle-ofattack rate. Therefore, longitudinal control manipulation by the pilot during the stall maneuver, close to the stall identification system activation 8-3 ### Page 166 05/04/18 AC 25-7D point, can advance or delay its activation without appreciably affecting the average stall entry airspeed rate. To minimize scatter in the stall speed versus entry rate data, the pilot should attempt to maintain a stable angleof-attack rate or pitch rate (not necessarily a fixed airspeed deceleration rate), until the stall identification system activates. The resulting timehistory of angle-of-attack data should be smooth and without discontinuities. A cross plot of airspeed deceleration rate, as defined in paragraph 8.1.4.5.5, versus angle-of-attack rate for all related test points, will show the general trend of this relationship for each flap setting. Any points that do not follow this general trend should not be used in establishing the stall speed. 8.1.4.4 Thrust Effects on Stall Speed. 8.1.4.4.1 Stall speeds are typically determined with the thrust levers at idle; however, it is necessary to verify by test or analysis that engine idle thrust does not result in appreciably lower stall speeds than would be obtained at zero thrust. Prior to amendment 25-108, a negative idle thrust at the stall, which slightly increases stall speeds, was considered acceptable, but applicants were not required to base stall speeds on idle thrust. With the adoption of amendment 25-108, it became a requirement to base stall speeds on idle thrust, except where that thrust level results in a significant decrease in stall speeds. If idle thrust results in a significant decrease in stall speeds, then stall speeds cannot be based on more than zero thrust. 8.1.4.4.2 To determine whether thrust effects on stall speed are significant, at least three stalls should be conducted at one flap setting, with thrust set to approximately the value required to maintain level flight at 1.5 VSR in the selected configuration. 8.1.4.4.3 These data may then be extrapolated to a zero thrust condition to determine the effects of idle thrust on stall speeds. (See figure 8-1 below.) If the difference between idle thrust and zero thrust stall speed is 0.5 knots or less, the effect may be considered insignificant. 8-4 ### Page 167 05/04/18 AC 25-7D Figure 8-1. Thrust Effect on Stall Speed **THRUST EFFECT ON** **STALL SPEED** **STALL SPEED ZERO THRUST THRUST FOR 1.5 VSR** **IN LEVEL FLIGHT IDLE THRUST** **ENGINE THRUST** 8.1.4.4.4 The effects of engine power on stall speeds for a turbopropeller airplane can be evaluated in a similar manner. Stall speed flight tests should be accomplished with engines idling and the propellers in the takeoff position. Engine torque, engine RPM, and estimated propeller efficiency can be used to predict the thrust associated with this configuration. 8.1.4.5 Data Reduction and Presentation. The following is an example of how the data obtained during the stall speed testing may be reduced to standard conditions. Other methods may be found acceptable. 8.1.4.5.1 Record the indicated airspeed from the flight test airspeed system throughout the stall and correct these values to equivalent airspeed. Also record load factor normal to the flight path. Typically, the load factor data would be obtained from a sufficient number of accelerometers capable of resolving the flight path load factor. It may be possible to obtain acceptable data using one accelerometer aligned along the expected 1-g stall pitch angle. More likely, it will take at least two accelerometers, one aligned along the fuselage longitudinal axis and one aligned at 90° to that axis, as well as a means to determine the angle between the flight path and the fuselage longitudinal axis. 8.1.4.5.2 Calculate the airplane lift coefficient (CL) from the equation given below and plot it as a time history throughout the stall maneuver: 8-5 ### Page 168 05/04/18 AC 25-7D Where: 8.1.4.5.3 The maximum lift coefficient CLMAX is defined as the maximum value of CL achieved during the stall test. Where the time history plot of CL exhibits multiple peak values, CLMAX normally corresponds to the first maximum. However, the peak corresponding to the highest CL achieved may be used for CLMAX, provided it represents usable lift, meaning that it does not occur after deterrent buffet or other stall identification cue (ref. § 25.201(d)). There should also typically be a noticeable break in a plot of the load factor normal to the flight path near the point at which CLMAX is reached. The analysis to determine CLMAXshould disregard any transient or dynamic increases in recorded load factor, such as might be generated by abrupt control inputs that do not reflect the lift capability of the airplane. The load factor normal to the flight path should be maintained at nominally 1.0 until CLMAX i s reached. (See figure 8-2 below.) 8-6 ### Page 169 05/04/18 AC 25-7D Figure 8-2. CLMAX and Load Factor 8.1.4.5.4 Where: 8-7 ### Page 170 8.1.4.5.5 Determine the stall entry rate, which is defined as the slope of a straight line connecting the stall speed and an airspeed 10 percent above the stall speed, for each stall test. Because CLMAX is relatively insensitive to stall entry rate, a rigorous investigation of entry rate effects should not be necessary. 8.1.4.5.6 For each approved configuration, construct a plot of CLMAX versus weight. (See figure 8-3 below.) Figure 8-3. CLMAX versus Weight and Flap Setting CL MAX 8.1.4.5.7 Flight test safety concerns usually dictate the lowest test altitude for determining stall speeds. The test data should then be expanded to lower altitudes, and hence lower Mach numbers, to cover the operational envelope of the airplane. Since CLMAX usually increases as the Mach number is reduced, simple expansion of the flight test data could result in extrapolating to a higher CL LMAX than tested. The expansion of C MAX versus Mach number data is only permitted up to the highest CLMAX demonstrated within the range of W/’s tested, unless the continuation of the trend of higher CLMAX with decreasing Mach number is substantiated with other test data. For example, data obtained at a more aft CG position or with power on can be used for this purpose if CG and thrust effects can be accounted for. Data from another airplane in the same family with the 8-8 ### Page 171 05/04/18 AC 25-7D same wing and showing the same general trend of CLMAX versus Mach (e.g., a lighter weight variant) may also be used if shown to be applicable. 8.1.4.5.8 The reference stall speed, VSR, is a calibrated airspeed defined by the applicant. VSR may not be less than the 1-g stall speed and is expressed as: Where: 8.1.4.5.9 Construct a plot of reference stall speed versus weight for each flap/gear configuration. (See figure 8-4 below). 8-9 ### Page 172 Figure 8-4. Stall Speed versus Weight and Flap Setting A B C D **INCREASING** **FLAP SETTING WEIGHT** **STALL SPEED (KEAS)** 8.1.4.5.10 For airplanes equipped with a device that abruptly pushes the nose down at a selected angle-of-attack (e.g., a stick pusher), VSR must not be less than the greater of 2 knots or 2 percent above the speed at which the device activates (§ 25.103(d)). 8.1.4.5.11 In showing compliance with § 25.103(d) for airplanes equipped a device that abruptly pushes the nose down at a selected angle-of-attack (e.g., a stick pusher), the speed at which the device operates need not be corrected to 1 g. Requiring a load factor correction of the device activation speed to the 1-g condition would unnecessarily increase the stringency of § 25.103(d). For example, it would be possible for the device activation speed to be assessed as higher than VSR (or at least closer to VSR than would be obtained without correcting to the 1 g condition). Test procedures should be in accordance with paragraph 8.1.4.3.1 to ensure that no abnormal or unusual pilot control input is used to obtain an artificially low speed at which the device first activates. 8.1.5 Stall Characteristics—§ 25.203. 8.1.5.1 Background. To assure a safe and expeditious recovery from an unintentional stall, it should not require any unusual piloting technique to successfully demonstrate compliance with § 25.203, nor should it require exceptional skill or repeated practice by the test pilot. The behavior of the airplane 8-10 ### Page 173 05/04/18 AC 25-7D during the stall and recovery must be easily controllable using normally expected pilot reactions. 8.1.5.2 Configuration. 8.1.5.2.1 Stall characteristics should be investigated with wings level and in a 30° banked turn, with both power or thrust on and power or thrust off in all configurations approved for normal operations. 8.1.5.2.2 The test configurations for stall characteristics should include deployed deceleration devices for all flap positions, unless limitations against the use of those devices with particular flap positions are imposed. Deceleration devices include spoilers used as airbrakes, and thrust reversers approved for inflight use. Stall demonstrations with deceleration devices deployed should normally be carried out with power or thrust off, except where deployment of the deceleration devices with power or thrust on would likely occur in normal operations (e.g., extended spoilers during landing approach). 8.1.5.2.3 Stall characteristics should be investigated with any systems or devices that may alter the stalling behavior of the airplane in their normal functioning mode. Unless the design of the airplane’s automatic flight control system precludes its ability to operate beyond the stall warning angle-of-attack, stall characteristics and the adequacy of stall warning should be evaluated when the airplane is stalled under the control of the automatic flight control system. 8.1.5.2.4 Power-off stalls should be conducted at flight idle for the appropriate configuration. For propeller-driven airplanes, the propeller should be set in the normal low pitch (high RPM) position. 8.1.5.2.5 For power-on stalls, power or thrust should be set to the value required to maintain level flight at a speed of 1.5 VSR at the maximum landing weight with flaps in the approach position, and the landing gear retracted. The approach flap position referred to is the maximum flap deflection used to show compliance with § 25.121(d), which specifies a configuration in which the reference stall speed does not exceed 110 percent of the reference stall speed for the related landing configuration. 8.1.5.2.6 Stall characteristics testing is normally done at the aft CG limit, which is typically the most adverse; however, if the stall speed tests at forward CG indicate that marginal stall recovery characteristics may exist at forward CG, compliance with § 25.203 should be shown for the most critical loading. 8.1.5.2.7 In accordance with § 25.21(c), stalls must be demonstrated up to the maximum approved operating altitude to determine if there are any 8-11 ### Page 174 adverse compressibility effects on stall characteristics. These tests should be flown with gear and flaps up at the most adverse CG. Power or thrust may be set, as required, to maintain approximately level flight and a 1 knot/second deceleration. A slight descent rate is permissible as long as the stall occurs at approximately the maximum approved altitude. Characteristics should be checked during a wings level stall and in a 30° banked turn. 8.1.5.2.8 For abnormal aerodynamic configurations covered by AFM procedures, high angle-of-attack characteristics should be evaluated down to the speed reached one second after stall warning in a one knot/second deceleration with the wings level and at idle power or thrust. If there are no adverse characteristics and there is adequate controllability, it is not necessary to stall the airplane. Adequate controllability means that it is possible to produce and to correct pitch, roll, and yaw by unreversed use of the flight controls, and that there are no uncommanded airplane motions due to aerodynamic flow breakdown. The applicant should also demonstrate that the airplane is safely controllable and maneuverable when flown at the recommended operating speed. 8.1.5.2.9 Stall characteristics should also be demonstrated with the maximum allowable asymmetric fuel loading. Requirements are as specified in § 25.203(a) and (c). 8.1.5.3 Procedures. 8.1.5.3.1 The airplane should be trimmed for hands-off flight at a speed 13 percent to 30 percent above the reference stall speed, with the appropriate power or thrust setting and configuration. Then, using only the primary longitudinal control, establish and maintain a deceleration (stall entry rate) consistent with that specified in § 25.201(c)(1) or (c)(2), as appropriate, until the airplane is stalled. Both power/thrust and pilot selectable trim should remain constant throughout the stall and recovery (to where the angle-of-attack has decreased to the point of no stall warning). 8.1.5.3.2 The same trim reference (for example, 1.23 VSR) should be used for both the stall speeds and characteristics testing. For all stall testing, the trim speed is based on the stall speeds provided in the AFM. 8.1.5.3.3 During the approach to the stall, the longitudinal control pull force should increase continuously as speed is reduced from the trimmed speed to the onset of stall warning. Below that speed some reduction in longitudinal control force is acceptable, provided it is not sudden or excessive. 8.1.5.3.4 Section 25.203(b) states that “the roll occurring between the stall and the completion of the recovery may not exceed approximately 20 degrees” for 8-12 ### Page 175 05/04/18 AC 25-7D level wing stalls. In level wing stalls the bank angle may exceed 20° occasionally, provided that lateral control is effective during recovery. 8.1.5.3.5 Section 25.203(c) requires the action of the airplane, following the 30° bank turning stalls, “not be so violent or extreme” such that a prompt recovery would be difficult and require more than normal piloting skill. The maximum bank angle that occurs during the recovery should not exceed approximately 60° in the original direction of the turn, or 30° in the opposite direction. 8.1.5.3.6 The intent of evaluating the 3 knot per second deceleration rate required under § 25.201(c)(2) is to demonstrate safe characteristics at higher rates of increase in angle-of-attack than are obtained from the 1 knot per second stalls. The specified airspeed deceleration rate, and associated angle-ofattack rate, should be maintained up to the point at which the airplane stalls. The maximum bank angle that occurs during the recovery should not exceed approximately 90° in the original direction of the turn, or 60° in the opposite direction. 8.1.5.3.7 For those airplanes where stall is defined by full nose-up longitudinal control for both forward and aft CG, the time at full aft stick during characteristics testing should be not less than that used for stall speed determination. For turning flight stalls, however, recovery may be initiated once the pitch control reaches the aft stop when accompanied by a rolling motion that is not immediately controllable (provided the rolling motion complies with § 25.203(c)). 8.1.5.3.8 As required by § 25.203(a), normal use of the lateral control must produce (or correct) a roll, and normal use of the directional control must produce (or correct) a yaw in the applied direction up to the point where the airplane is considered stalled. It must be possible to prevent or recover from a stall by normal use of the controls. 8.1.5.3.9 If wind tunnel tests have indicated an airplane may be susceptible to deep stall penetration (i.e., that area beyond the stall angle-of-attack from which recovery may be difficult or impossible), substantiation should be provided that there is adequate recovery control available at, and sufficiently beyond, the stall angle-of-attack. 8.1.6 Stall Warning—§ 25.207. 8.1.6.1 Explanation. The purpose of these stall warning requirements is to provide an adequate spread between warning and stall to allow the pilot time to recover without inadvertently stalling the airplane. 8-13 ### Page 176 05/04/18 AC 25-7D 8.1.6.2 Background. To be acceptable, a stall warning must have the following features: 8.1.6.2.1 Distinctiveness. The stall warning indication must be clear and distinct to a degree that will ensure positive pilot recognition of an impending stall. 8.1.6.2.2 Timeliness. For one knot per second entry rate stalls, the stall warning must begin at a speed, VSW, not less than five knots or five percent (whichever is greater) above the speed at which the stall is identified in accordance with § 25.201(d). For straight flight stalls, at idle power or thrust and with the CG at the position specified in § 25.103(b)(5), the stall warning must begin at a speed not less than three knots or three percent (whichever is greater) above the reference stall speed. These speed margins should be in terms of the same units of measurement as VSR (i.e., calibrated airspeed). 8.1.6.2.3 Consistency. The stall warning should be reliable and repeatable. The warning must occur with flaps and gear in all normally used positions in both straight and turning flight (§ 25.207(a)) and must continue throughout the stall demonstration until the angle-of-attack is reduced to approximately that at which the stall warning was initiated (§ 25.207(c)). The warning may be furnished naturally through the inherent aerodynamic characteristics of the airplane, or artificially by a system designed for this purpose. If artificial stall warning is provided for any airplane configuration, it must be provided for all configurations (§ 25.207(b)). 8.1.6.2.4 An artificial stall warning indication that is a solely visual device which requires attention in the cockpit, inhibits cockpit conversation or, in the event of malfunction, causes distraction that would interfere with safe operation of the airplane, is not acceptable. 8.1.6.2.5 For airplanes that use artificial stall warning systems, paragraph 42.1 of this AC presents guidance material for demonstrating compliance with the regulatory requirements of part 25. 8.1.6.2.6 If the stall warning required by § 25.207 is provided by an artificial stall warning system (e.g., a stick shaker), the effect of production tolerances on the stall warning system should be considered when evaluating the stall warning margin required by § 25.207(c) through (f) and the maneuver capabilities required by § 25.143(h). 8.1.6.2.7 The stall warning margin required by § 25.207(c) through (f) should be available with the stall warning system set to the most critical setting expected in production. Unless another setting would provide a lesser 8-14 ### Page 177 05/04/18 AC 25-7D margin, the stall warning system should be operating at its high angle-ofattack limit. For airplanes where VSR is set by a device that abruptly pushes the nose down at a selected angle-of-attack (e.g., a stick pusher), the stall warning margin may be evaluated with both the stall warning and stall identification (e.g., stick pusher) systems at their nominal angle-ofattack settings unless a lesser margin can result from the various system tolerances. 8.1.6.2.8 The maneuver capabilities required by § 25.143(h) should be available assuming the stall warning system is operating on its nominal setting. When the stall warning system is operating at its low angle-of-attack limit, the maneuver capabilities should not be reduced by more than 2° of bank angle from those specified in § 25.143(h). A flight test, an acceptable analysis, or simulation can be used to make this assessment. 8.1.6.2.9 The stall warning margin and maneuver capabilities may be demonstrated by flight testing at the most critical settings specified above for the stall warning and, if so equipped, stall identification systems. Alternatively, compliance may be shown by applying adjustments to flight test data obtained at a different system setting if an acceptable method is used that takes into account all of the relevant variables. 8.1.6.3 Procedures. Stall warning tests are normally conducted in conjunction with the stall testing required by § 25.103 (stall speeds), § 25.201 (stall demonstration), and § 25.203 (stall characteristics), including consideration of the prescribed bank angles, power or thrust settings, and CG position. The pilot technique in stalling the airplane should be consistent between the onset of stall warning and the point at which the stall is identified. That is, there should not be any deliberate attempt to reduce the load factor, change the deceleration, or use any other means to increase the stall warning margin. In addition, if the stall warning margin may be affected by a system (e.g., a stall warning or stick pusher system that modifies the stall warning or stall identification speed as a function of power or thrust, bank angle, angle-of-attack rate, etc.), compliance with § 25.207(c) should be demonstrated at the most critical conditions in terms of stall warning margin. However, for this case, bank angles greater than 40° and power or thrust exceeding maximum continuous power or thrust need not be demonstrated. If the effect of the stall identification or stall warning system compensation is to increase the stall warning margin relative to the nominal values demonstrated during the testing required by §§ 25.103, 25.201, and 25.203, these additional stall warning margin demonstrations need not be done. 8-15 ### Page 178 8.1.6.4 Data Acquisition and Reduction. The stall warning speed and type and quality of warning should be noted. To determine if the required margin exits, compare the speed at which acceptable stall warning begins with (1) the stall identification speed, and (2) VSR (for the conditions under which VSR is defined). The stall warning margin comparisons should be made at a constant 1-g load factor when showing compliance with § 25.207(d). 8.1.7 Accelerated Stall Warning. 8.1.7.1 Explanation. Section 25.207(f) requires that, in slow-down turns with at least a 1.5 g load factor normal to the flight path and an airspeed deceleration rate greater than 2 knots per second, sufficient stall warning is provided to prevent stalling when the pilot takes recovery action not less than one second after recognition of stall warning. The purpose of the requirement is to ensure that adequate stall warning exists to prevent an inadvertent stall under the most demanding conditions that are likely to occur in normal flight. The elevated load factor will emphasize any adverse stall characteristics, such as wing drop or asymmetric wing flow breakdown, while also investigating Mach and potential aeroelastic effects on available lift. 8.1.7.2 Procedures. 8.1.7.2.1 Trim at 1.3 VSR. Once trimmed, accelerate to a speed that will allow enough time to set up and complete the maneuver at the specified load factor and airspeed deceleration rate. Set power or thrust appropriate to the power or thrust for level flight at 1.3 VSR and do not adjust it during the maneuver. In a level flight maneuver, 1.5 g equates to a bank angle of 48°. To prevent an excessive deceleration rate (e.g., greater than 3 knots per second), a descent may be used. Conversely, if the deceleration rate is too low, the maneuver should be conducted in a climbing turn. 8.1.7.2.2 After the onset of stall warning, continue the maneuver without releasing stick force for one second before attempting recovery. Normal low speed recovery techniques should be used. If any of the indications of a stall prescribed in § 25.201(d) (see paragraph 8.1.3.1 of this AC) occur during the accelerated stall warning demonstration, compliance with § 25.207(f) will not have been demonstrated. 8.1.8 Maneuver Margins. See paragraph 5.1 of this AC for guidance material associated with demonstrating compliance to the maneuvering capability requirements of § 25.143(h). 8-16 ### Page 179 8.1.9 Additional Considerations for Airplanes Equipped with Stall Identification Systems. A stall identification system is any system that is used to show compliance with § 25.201(d), which requires the airplane to give the pilot a clear and distinctive indication of stall. The stall identification system consists of everything from the sensing devices that supply inputs to the system to the activation of the system response that provides stall identification to the flightcrew. Section 25.1309(a) requires that such a system, when it is needed to show compliance with the stall-related requirements, be designed to perform its intended function under any foreseeable operating condition. 8.1.9.1 The applicant should verify that the stall identification system, considering system design features (e.g., filtering, phase advancing) and airplane and system production tolerances will not result in an unsafe diminishing of the margin between stall warning and stall identification, or between stall identification and any hazardous airplane flight characteristic in any foreseeable operating condition. This verification may be provided by a combination of analysis, simulation, and flight test. The following operating conditions should not result in unwanted activation of the stall identification system or in aerodynamic stall prior to, or close to, activation of the stall warning system: dynamic and accelerated stall entries, the effects of atmospheric turbulence, any foreseeable type of wing contamination (e.g., ice, frost, insects, dirt, anti-icing fluids), or wing leading edge damage within prescribed maintenance limits. Operation in windshear environments where the airplane will be flown at, or very near, stall warning, should also be considered, although, depending on the severity of the windshear, it may be impossible to ensure that there is no possibility of stall indication system operation. For wing contamination, the applicant should substantiate the critical height and density of the contaminant. Carborundum sandpaper No. 40 (that is, 40-grit carborundum sandpaper) has been used in past certification programs to represent residual ice or frost contamination. 8.1.9.2 Stall characteristics testing should be performed with the following airplane and stall identification system production tolerances set to achieve the most adverse stall identification system activation condition for stall characteristics: 8.1.9.2.1 Airframe build tolerances—the impact of wing angle of incidence variation relative to stall identification system vane angle; and 8.1.9.2.2 Stall identification system tolerances (e.g., activation vane angles). 8.1.9.3 If the combined root-sum-square (square root of the sum of the squares of each tolerance) effect of the tolerances identified above is less than ±1 knot, stall speeds testing can be performed and the stall speeds determined with the tolerances at their nominal values. If the combined root-sum-square effect is ±1 knot or greater, stall speed testing should be 8-17 ### Page 180 05/04/18 AC 25-7D performed with the tolerances at the values that would result in the highest stall speeds. 8.1.10 Reliability of Artificial Stall Warning and Stall Identification Systems. Additional guidance material related to the testing and approval of artificial stall warning and stall identification systems is presented in paragraph 42.1 of this AC. 8-18 ### Page 181 CHAPTER 9. FLIGHT: GROUND AND WATER HANDLING CHARACTERISTICS 9.1 Part 25 Regulations. The applicable regulations are §§ 25.231, 25.233, 25.235, 25.237, and 25.239. 9.2 Longitudinal Stability and Control—§ 25.231. 9.2.1 Explanation. Test program objectives would not be expected to demonstrate taxiing over rough surfaces at speeds high enough to approach structural design limits, nor is it expected that in the test program the airplane be landed harder or at higher sink rates than it will ever encounter in service. However, new or modified landing gear systems should be evaluated on rough surfaces that are representative of normal service, and landings should be conducted at various sink rates sufficient to identify any dangerous characteristics or tendencies. Variables to be considered are CG and taxi speed. The cockpit motion dynamics during ground handling should not impede control of the airplane, and pitching motion during bounce should not create static pitch control problems or pilot induced oscillation tendencies. 9.2.2 Procedures. Ground handling tests at speeds normally expected in service should be conducted on smooth and rough surfaces that are likely to be encountered under normal operating conditions. Particular attention should be paid to the following: 9.2.2.1 Brakes. The adequacy of the brakes when maneuvering on the ground and the tendency of the brakes to cause nosing-over should be investigated. Any bad tendency will normally be exaggerated when taxiing in a strong cross or tail wind. 9.2.2.2 Seaplanes and Amphibians. The most adverse water conditions safe for taxiing, takeoff, and landing must be established per § 25.231(b). Procedure and limitations for using reverse thrust should be determined. 9.3 Directional Stability and Control—§ 25.233. 9.3.1 Explanation. None. 9.3.2 Procedures. Taxi, takeoff, and landing should be conducted in all configurations under normal operating conditions. 9-1 ### Page 182 **AC 25-7D CHG 1** 9.3.2.1 There may be no uncontrollable ground-looping tendency in 90° crosswinds, up to an averaged wind velocity of 20 knots or 0.2 VSR0, whichever is greater (except that the wind velocity need not exceed 25 knots) at any speed at which the airplane may be expected to be operated on the ground. This may be shown while establishing the 90° crosswind component required by § 25.237. 9.3.2.2 Landplanes must be satisfactorily controllable, without exceptional piloting skill or alertness in power-off landings at normal landing speed, without using brakes or engine power or thrust to maintain a straight path. This may be shown during power-off landings made in conjunction with other tests. 9.3.2.3 The airplane must have adequate directional control during taxiing. This may be shown during taxiing prior to takeoffs made in conjunction with other tests. 9.4 Taxiing Condition—§ 25.235. [Reserved] 9.5 Wind Velocities—§ 25.237. 9.5.1 Explanation. 9.5.1.1 Landplanes. 9.5.1.1.1 There must be a 90° crosswind component established that is shown to be safe for takeoff and landing on dry runways. 9.5.1.1.2 The airplane must exhibit satisfactory controllability and handling characteristics in 90° crosswinds at any ground speed at which the airplane is expected to operate. 9.5.1.2 Seaplanes and Amphibians. 9.5.1.2.1 There must be a 90° crosswind component established that is shown to be safe for takeoff and landing in all water conditions that may reasonably be expected in normal operation. 9.5.1.2.2 There must be a wind velocity established for which taxiing is safe in any direction under all water conditions that may reasonably be expected in normal operation. 9-2 ### Page 183 **AC 25-7D CHG 1** 9.5.1.3 Crosswind Demonstration. An averaged 90° crosswind component at 10 meters (as required by § 25.21(f)) of at least 20 knots or 0.2 VSR0 (where VSR0 is for the maximum design landing weight), whichever is greater, except that it need not exceed 25 knots, must be demonstrated during type certification tests. At the same time, the maximum gusts encountered should be established and their effect on airplane handling characteristics in crosswind assessed. There are two results possible: 9.5.1.3.1 A crosswind component value may be established that meets the minimum requirements but is not considered to be a limiting value for airplane handling characteristics. This demonstrated value should be included as information in the AFM. 9.5.1.3.2 A crosswind component value may be established that is considered to be a maximum limiting value up to which it is safe to operate for takeoff and landing. This limiting value should be shown in the operating limitations section of the AFM. Note: The crosswind component included in AFM, whether limiting or not, should be provided as a single gust included value i.e. “XX kt (Gust included)”. A set of two values, such as “Average XX kt with gusts up to YY kt”, is acceptable, although not preferred. Other formats, in particular those not providing information related to gusts, should not be used. 9.5.2 Procedures. 9.5.2.1 Configuration. These tests should be conducted in the following configurations: 9.5.2.1.1 At light weight and aft CG. (This is desirable; however, flexibility should be permitted.) 9.5.2.1.2 Normal takeoff and landing flap configurations using the recommended procedures. 9.5.2.1.3 Normal usage of thrust reversers. Particular attention should be paid to any degradation of rudder effectiveness due to thrust reverser airflow effects. 9.5.2.1.4 Yaw dampers/turn coordinator On, or Off, whichever is applicable. 9-3 ### Page 184 **AC 25-7D CHG 1** 9.5.2.2 Test Procedures. Three takeoffs and three landings, with at least one landing to a full stop, should be conducted in a 90° crosswind component of at least 20 knots or 0.2 VSR0, whichever is greater, except that it need not exceed 25 knots. For each test condition, a qualitative evaluation by the pilot of airplane control capability, forces, airplane dynamic reaction in gusty crosswinds (if available), and general handling characteristics should be conducted. The airplane should be satisfactorily controllable without requiring exceptional piloting skill or strength. If thrust reversers are installed, these landings should be conducted with the thrust reversers deployed as per normal procedures and additional landings should be conducted at the critical reverse thrust/power level to verify that there are no unsatisfactory handling characteristics. 9.5.2.3 Test data. Crosswind data may be obtained from a calibrated flight test wind measurement station, from an airfield wind reporting device, or from any other method acceptable to the FAA. 9.5.2.3.1 A calibrated flight test wind measurement station located in the vicinity of the liftoff or touchdown point generally provides the most accurate data and is preferable. 9.5.2.3.2 An airport wind reporting device may also be acceptable provided the device has been calibrated and is located near the runway being used for testing. 9.5.2.3.3 Crosswind data taken directly from a commercially available inertial or differential GPS based reference system may not be accurate in sideslips and is not accurate on the ground. During landing, filtering may introduce lags making the data incorrect due to wind shear with altitude (i.e., a higher wind value at altitude is “remembered”). Hence this method is considered unsuitable for accurately determining the crosswind during takeoff and landing. 9.5.2.3.4 Other methods based on the computation of the actual crosswind encountered by the airplane based on on-board measurements are also acceptable. For example, the crosswind can be computed by resolving the difference between true airspeed (from an ADC) and an accurate ground speed measurement (e.g., derived from IRS groundspeed) into the alongrunway and across-runway heading taking into account the airplane heading, track angle and sideslip. 9-4 ### Page 185 **AC 25-7D CHG 1** 9.5.2.3.5 No matter which method is used, the wind should be continuously timerecorded throughout the takeoff from brake release (or any low speed above which all data necessary to the computation are available and of sufficient accuracy) to a height of 50 feet, and throughout the landing from a height of 50 feet to termination of the test event (e.g., full stop, touchand-go, go-around) or any low speed above which all data necessary to the computation are available and of sufficient accuracy. The measured crosswind component should be corrected from the height of the measurement device to a height of 10 meters. The average crosswind at 90° to the runway heading should then be calculated for the above time span. The maximum gust could also be derived during this process, based on the same time span. 9.5.2.3.6 With prior agreement from the FAA, it may also be permissible to obtain crosswind data from tower wind reports. However, the use of this method should be carefully reviewed to ensure that the measurement sensor is properly calibrated to establish the measurement sensor reference height, to establish that the smoothing characteristics do not produce unacceptable filtering, and that the location of the measurement sensor is appropriate for the takeoff and landing runway(s). Such a method has the disadvantage of not being able to provide the gust value during takeoff and landing. 9.5.2.3.7 For all methods of gathering test data (with the exception of the method described in paragraph 9.5.2.3.6), the following is applicable: The averaged value of the crosswind should be understood as a mathematical average obtained from: Where Vwy is the crosswind component corrected to 10m height, per § 25.21(f), and perpendicular to the runway axis. 9-5 ### Page 186 **AC 25-7D CHG 1** 1. The maximum gust derived from test analysis should be of a duration sufficient to interfere with airplane handling characteristics in crosswind. It should be obtained from a centered moving average applied to the test data. The centered moving average should not be of less than 3 seconds, in order to be consistent with the filtering standard applied by airports to communicate the gust value. Note, however, that a 3 second moving average applied on the data collected during the tests may not necessarily provide the same result as airport reported gusts, due to differences in data acquisition and reduction methodologies between flight tests systems and airport systems. This filtering also introduces some conservatism in the gust determination to account for the variability of gusts profiles, as compared to the ones encountered in flight tests. 2. If an applicant wants to use a centered moving average below 3 seconds, this should be substantiated on a case-by-case basis. The moving average should not go below 1 second. 3. Using a smaller portion of the takeoff or landing run (i.e. computing the average starting below 50 ft and/or above the first/last valid data) or only relevant portions of the takeoff or landing run to compute the average crosswind may be acceptable if properly justified by the applicant. 9.5.2.3.8 If demonstrated crosswind is not considered limiting, an applicant may use a separate and distinctively identified portion of the AFM to provide “unapproved” data, regarding airplane crosswind capability beyond the demonstrated level. 9.6 Spray Characteristics, Control, and Stability on Water—§ 25.239. 9.6.1 Explanation. These characteristics should be investigated at the most adverse weight/CG combinations. 9.6.2 Procedures. 9.6.2.1 The spray characteristics and, in particular, the pilot view during the initial takeoff run, should allow sufficient view in order to maintain a reasonable track over the water. Since not all seaplane operations are on open lakes or bays, but can be on rivers or channels, the directional control and view should be sufficient enough to stay within the channel confines. 9.6.2.2 The tendency of the wing floats or sponsons to submerge and/or cause waterloops should be evaluated during the crosswind testing. During the step taxiing evaluations, the floats should also be evaluated for any 9-6 ### Page 187 AC 25-7D CHG 1 tendency to bury and either cause waterlooping or damage. The procedures used to avoid undesirable characteristics should be included in the AFM. 9.6.2.3 During low-speed taxi, the effectiveness of the water rudders and/or asymmetric power or thrust should be evaluated in view of the types of maneuvering to be expected in service. If reverse thrust is to be used, it too should be evaluated in terms of ease of accomplishment and crew coordination. 9.6.2.4 If an amphibian is intended to be “beached” or run up a ramp, the handling characteristics and ability to maneuver onto the ramp should be evaluated. Forward CG is generally more critical. The procedures should be included in the AFM. There should be no undue tendency to damage the bow or other structure. 9.6.2.5 Engine failure of the critical engine at any time during the takeoff run should be evaluated. No dangerous porpoising, swerving, or waterlooping should result. 9.6.2.6 There should be no undue tendency to porpoise and no extraordinary skill or alertness should be required to control porpoising. 9.6.2.7 Spray impingement on the airframe (control surfaces, etc.) should be evaluated to assure the resulting loads are within acceptable limits. 9.6.2.8 The above evaluations should be performed in the airplane on the water rather than by analysis or model testing. Analysis and/or model testing may be used to point out the problem areas but should not be substituted for actual testing. 9-7 ### Page 188 CHAPTER 10. FLIGHT: MISCELLANEOUS REQUIREMENTS 10.1 Vibration and Buffeting—§ 25.251. 10.1.1 Explanation. 10.1.1.1 The testing required by subpart C of part 25 covers the vibration extremes expected in service. The applicant’s flight tests should assure that the regulatory limits are not exceeded. Flight testing should not be conducted beyond where structural (subpart C) tests and calculations have been completed. 10.1.1.2 For § 25.251(b) and (c), vibration and buffeting are considered excessive when it is determined that it: 10.1.1.2.1 May cause structural damage or, if sustained over an extended period of time, could lead to structural fatigue; 10.1.1.2.2 May cause pilot fatigue or annoyance that interferes with operation of the airplane or management of the airplane systems; or 10.1.1.2.3 Interferes with flight instrument readability. 10.1.1.3 No perceptible buffeting is permitted in the cruise configuration as required by § 25.251(d). Weight and/or altitude AFM limitations may need to be imposed to comply with this criterion. Reasonable buffet during the deployment of spoilers and other high drag devices is permitted to the extent allowed under § 25.251(b) and (c), as described in paragraph 10.1.1.2 above. 10.1.1.4 For airplanes with MD greater than 0.6 or with a maximum operating altitude greater than 25,000 feet, the buffet onset envelope must be established for the ranges of airspeed and/or Mach number, weight, altitude, and load factor for which the airplane is to be certificated. This envelope must be provided in the AFM in accordance with § 25.1585(d). These AFM data should be valid criteria for forward CG conditions or correctable to forward CG using AFM procedures. This boundary should be established by pilot qualitative evaluation or by correlation with pilot qualitative evaluation, as there is no predetermined criterion for buffet level at the pilot station. A normal acceleration of ±0.05 g has been used in some cases; however, the appropriate acceleration level will vary from airplane to airplane and may also be affected by the dynamic response of the accelerometer. If a measured normal acceleration is to be used, the acceleration level and specific accelerometer should first be correlated against a pilot’s assessment of the onset of buffet. 10-1 ### Page 189 10.1.1.5 Modifications to airplanes, particularly modifications that may affect airflow about the wing, should be evaluated for their effect on vibration and buffeting characteristics, changes in the speeds for onset of buffet, and maneuvering characteristics beyond buffet onset. This change may not only impact the buffet boundary envelope but may change the acceptability of the VMO/MMO or VDF/MDF speeds established on the unmodified airplane. If this occurs, the maximum operating speed and demonstrated flight diving speed may need to be reduced. However, the regulations concerning the speed spread margin between VMOMMO and VDF/MDF remain in effect. Systems and flight characteristics affected by the reduced maximum speeds should also be reevaluated. Indicator markings, overspeed horns, etc. must be reset, as necessary, to remain in compliance with the applicable regulations. 10.1.1.6 On swept-wing airplanes, undesirable pitch-up maneuvering characteristics can occur as the center of lift moves inboard and forward with increasing g, due to shockwave-induced separation and/or as wing load alleviation systems unload the wingtips. Straight-wing airplanes can also exhibit similar characteristics; therefore, new airplanes and those modified in a manner that may affect the spanwise lift distribution or produce undesirable pitching moment as a function of g, or increase the exposure to high altitude buffet encounters, should be evaluated as described herein. 10.1.1.7 Section 25.251(e) requires that “probable inadvertent excursions beyond the boundaries of buffet” may not result in “unsafe conditions.” To assure that no unsafe conditions are encountered in maneuvering flight, maneuvering flight evaluations to demonstrate satisfactory maneuvering stability are described below. A determination of the longitudinal maneuvering characteristics should be made to assure the airplane is safely controllable and maneuverable in the cruise configuration to assure there is no danger of exceeding the airplane limit load factor, and that the airplane’s pitch response to the primary longitudinal control is predictable to the pilot. 10.1.2 Procedures. 10.1.2.1 Section 25.251(a). The test procedures outlined below will provide the necessary flight demonstrations for compliance with § 25.251(a). 10.1.2.2 Section 25.251(b). The airplane should be flown at VDF/MDF at several altitudes from the highest practicable cruise altitude to the lowest practicable altitude. The test should be flown starting from trimmed flight at VMO/MMO at a power or thrust setting not exceeding maximum continuous power or thrust. The airplane gross weight should be as high as practicable for the cruise 10-2 ### Page 190 AC 25-7D CHG 1 condition, with the CG at or near the forward limit. In addition, compliance with § 25.251(b) should be demonstrated with high drag devices (i.e., speed brakes) deployed at VDF/MDF. Thrust reversers, if designed for inflight deployment, should be deployed at their limit speed conditions. 10.1.2.3 Section 25.251(c). The weight of the airplane should be as heavy as practical, commensurate with achieving the maximum certificated altitude. 10.1.2.4 Section 25.251(d). It should be demonstrated in flight tests that perceptible buffeting does not occur in straight flight in the cruise configuration, at any speed up to VMO/MMO, to show compliance with § 25.251(d). This should be met from initial combinations of critical weight and altitude, if achievable, where the airplane has a 0.3 g margin to the buffet onset boundary developed under § 25.251(e). These initial conditions should be established using a nominal cruise Mach number (typically long-range cruise Mach, MLRC) with the CG at the forward limit. This flight condition is representative of practical operating criteria imposed by most operators. From these initial conditions, the airplane should be accelerated in 1 g flight to VMO/MMO using maximum continuous power or thrust. Descending flight is acceptable if needed to achieve VMO/MMO. 10.1.2.5 Section 25.251(e). Section 25.251(e) requires the determination of the buffet onset envelope, in the cruise configuration, for airplanes with MD greater than 0.6 or maximum operating altitudes greater than 25,000 feet. This requirement also provides criteria for evaluation of maneuvering stability in cruise flight under load factor conditions up to and beyond the onset of buffet. 10.1.2.5.1 The determination of compliance with § 25.251(e), using flight test data from maneuvers conducted well into buffet, is extremely difficult due to the dynamics of this type of maneuver and the establishment of the FS/g relationship from such data. The pilot flying the airplane needs to evaluate the airplane characteristics under such conditions. Figure 10-1 provides guidance on stick force per g (FS/g) characteristics that would be considered acceptable or unacceptable. 10.1.2.5.2 For determination of the buffet onset envelope, the flight tests should be conducted at forward CG. For maneuvering characteristics, airplanes should be evaluated at the most aft CG in accordance with the following criteria: 4. For all weight/altitude combinations where buffet onset occurs at various load factors between approximately +1 g and +2 g, the longitudinal control force (FS) vs. g curve should have a positive slope 10-3 The text continues at https://openairdocs.org/ac/25-7.md?from=191 --- OpenAirDocs.org is an independent research project of [Contrail Labs](https://contraillabs.com/). It is not affiliated with the FAA or the NTSB. Source documents are U.S. government works in the public domain.