# AC 27-1B CHG 1-2: Certification of Normal Category Rotorcraft - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Historical - Version in force: [AC 27-1B CHG 1-9](https://openairdocs.org/ac/27-1), approved June 23, 2023. This is a historical version - Approved: April 25, 2006 - Last reviewed: December 4, 2008 - Family: AC 27-1 - 14 CFR parts: 27 - Pages: 932 - PDF: https://openairdocs.org/files/sha256/d7/66/d76666c472d0836e475634db9b0a8f57fef1da148eb89fbd871791b3a4faa26d - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/8C8DDE08737322428625719D0066C561.0001 - DRS record last changed: November 22, 2025 - Page: https://openairdocs.org/ac/27-1B-CHG-1-2 ## Versions - [AC 27-1B CHG 1-9](https://openairdocs.org/ac/27-1) (Current), approved Jun 23, 2023, 1,196 pages - [AC 27-1B](https://openairdocs.org/ac/27-1B?v=DRSDOCID123898489720230413220327) (Historical), approved Jun 23, 2023, 1,196 pages - [AC 27-1B CHG 1-8](https://openairdocs.org/ac/27-1B-CHG-1-8) (Historical), approved Jun 29, 2018, 1,140 pages - [AC 27-1B CHG 1-2](https://openairdocs.org/ac/27-1B-CHG-1-2) (Historical), approved Apr 25, 2006, 932 pages - [AC 27-1B](https://openairdocs.org/ac/27-1B?v=0CE1B2A7CEE21AAE8625703800684002) (Historical), approved Jun 24, 2005, 7 pages. CHGS to Various Paragraphs - [AC 27-1B](https://openairdocs.org/ac/27-1B?v=771532252D914FF686256D42006468E0) (Historical), approved Apr 25, 2003, 3 pages - [AC 27-1B](https://openairdocs.org/ac/27-1B?v=897391F05A323C6186256CEE005D87AF) (Historical), approved Feb 12, 2003, 2 pages - [AC 27-1B](https://openairdocs.org/ac/27-1B?v=6F95F1C7B8B1E89C86256E4300597919) (Historical), approved Feb 12, 2003, 901 pages - [AC 27-1A](https://openairdocs.org/ac/27-1A) (Historical), approved Jul 30, 1997, 690 pages - [AC 27-1B](https://openairdocs.org/ac/27-1B?v=5232FE84F9AF806C8625707B007183F6) (Historical), approved , 4 pages. 27.351 - Yawing Conditions ## ADs that cite it (1) - [AD 2018-02-01](https://openairdocs.org/ad/2018-02-01): Main Rotor Head · Enstrom Helicopter Corporation, The 280, 280C, 280F and 8 more. Effective Feb 21, 2018 ## Text Each page's text from the PDF; open the PDF at a page with #page=N. Some pages are OCR from the scanned FAA advisory circular. OCR can misread letters and numbers. Those pages are marked (OCR). ### Page 1 **TIPS ON USING BOOKMARKS IN THIS** **DOCUMENT** We have added a new feature that should help you to move around more quickly in the advisory circular. On the upper left side, click on “Bookmarks.” You can click on any line that has a plus sign (+) beside it and it will expand to show you everything that comes under that particular section. You can click on any bookmark and you will go directly to that paragraph. ### Page 2 Subject: CERTIFICATION OF NORMAL CATEGORY ROTORCRAFT Date: Initiated 4/25/06 ASW-110 AC No: Change 27-1B 2 ______________________________________________ **1. PURPOSE.** a. This change incorporates all the previously revised AC paragraphs that were posted as accepted and finalized AC material on the FAA RGL website since 2/12/03. b. This change revises existing material in 16 paragraphs and adds new material for one paragraph. c. The change number and the date of the changed material are shown at the top of each page. The vertical lines in the right and left margins indicate the beginning and end of each change. Pages that have different page numbers but no text changes will retain the previous heading information. d. This AC does not change regulatory requirements and does not authorize changes in, or deviations from, regulatory requirements. This AC establishes an acceptable means, but not the only means, of compliance. Since the guidance material presented in this AC is not regulatory, terms having a mandatory definition, such as “shall” and “must,” etc., as used in this AC, apply either to the reiteration of a regulation itself, or to an applicant who chooses to follow a prescribed method of compliance without deviation. **2. PRINCIPAL CHANGES.** a. Paragraphs 27.351, 27.602, 27.672, 27.683, 27.777, 27.1321, 27.1585, MG 1, MG 4, MG 8, MG 16, Appendix A, and Appendix B are revised. b. The AC material in paragraphs 27.801 and 27.1411 has been revised. c. The AC material in MG 12 has been revised and is now contained in AC 27.865B, in Subpart D. d. New paragraph MG 18, Helicopter Terrai n and Awareness Systems (HTAWS), is added to Chapter 3. e. New figures AC 27.351-2 and AC 27.865B-1 are added in Chapter 2. f. New figures AC 27 MG 8-1, MG 8-2, MG 8-3, MG 8-4, MG 18-1, MG 18-2, and MG 18- 2.1 are added in Chapter 3. 3. WEBSITE AVAILABILITY. To access this AC electronically, log on to http://www.airweb.faa.gov/rgl and then click on AC’s. ### Page 3 AC 27-1B, Chg 2 4/25/06 AC 27-1B, Chg 2 PAGE CONTROL CHART Remove pages Dated Insert Pages Dated v thru xxi........................... 9/30/99 & 2/12/03 C-9 thru C-86.................... 9/30/99 & 2/12/03 D-1 thru D-58 ................... 9/30/99 & 2/12/03 D-67 thru D-70 ...................................9/30/99 D-83 thru D-112 ............... 9/30/99 & 2/12/03 F-33 thru F-102...................................9/30/99 G-43 thru G-48 ...................................9/30/99 MG 1-15 thru MG 1-17 ......................9/30/99 MG 4-9 thru MG 4-13 ........................9/30/99 MG 8-1 thru MG 8-15 ...... 9/30/99 & 2/12/03 MG 12-1 thru MG 12-42 .. 9/30/99 & 2/12/03 MG 16-1 thru MG 16-16 ....................2/12/03 none none Apdx A-47 thru Apdx A-70................2/12/03 Apdx B-11 thru Apdx B-20 ................9/30/99 v thru xxii ...........................................4/25/06 C-9 thru C-87......................................4/25/06 D-1 thru D-58 .....................................4/25/06 D-67 thru D-70 ...................................4/25/06 D-83 thru D-134 .................................4/25/06 F-33 thru F-103 ..................................4/25/06 G-43 thru G-49 ...................................4/25/06 MG 1-15 thru MG 1-16 ......................4/25/06 MG 4-9 thru MG 4-13 ........................4/25/06 MG 8-1 thru MG 8-31 ........................4/25/06 MG 12-1 .............................................4/25/06 MG 16-1 thru MG 16-18 ....................4/25/06 MG 17-1 .............................................4/25/06 MG 18-1 thru MG 18-27 ....................4/25/06 Apdx A-47 thru Apdx A-71 ...............4/25/06 Apdx B-11 thru Apdx B-20................4/25/06 David A. Downey Manager, Rotorcraft Directorate Aircraft Certification Service ### Page 4 Subject: CERTIFICATION OF NORMAL CATEGORY ROTORCRAFT Date: Initiated 2/12/03 ASW-110 AC No: Change 27-1B 1 ______________________________________________ **1. PURPOSE.** a. This change revises existing material in 17 paragraphs and adds new material for 12 paragraphs. b. The change number and the date of the changed material are shown at the top of each page. The vertical lines in the right and left margins indicate the beginning and end of each change. Pages that have different page numbers but no text changes will retain the previous heading information. **2. CANCELLATION.** a. AC 20-95, Fatigue Evaluation of Rotorcraft Structure, May 18, 1976, is canceled in its entirety, and is replaced by material contained in AC 27 MG 11. b. AC 20-137, Dynamic Evaluation of Seat Restraint Systems & Occupant Restraint for Rotorcraft (Normal and Transport), March 30, 1992, is canceled in its entirety, and is replaced by material contained in AC 27.562. c. AC text containing references to AC 20-95 and AC 20-137 was either changed or deleted in paragraphs AC 27.2, 27.562, 27.613, 27.785A, 27.865A, 27.907, 27.952, MG 8, and MG 12. **3. PRINCIPAL CHANGES.** a. Paragraphs 27.562, 27.571, 27.1303, 27.1353, 27.1419, 27.1505, 27.1529, MG 5, MG 6, and MG 12 are revised. b. New paragraphs 27.610A, 27.625A, 27. 785B, 27.805, 27.807B, and 27.853A are added to Chapter 2. c. New paragraphs MG 11, MG 13, MG 14, MG 15, and MG 16 are added to Chapter 3. d. Appendix A (previously reserved) is added. e. New figures AC 27.562-1, 27.562-2, 27.562-3, 27.562-4, 27.562-5, 27.562-6, 27.562-7, and 27.562-8 are added in Chapter 2. f. New figures AC 27 MG 11-1, 11-2, 11-3, 11-4, 11-5, 11-6, 11-7, 11-8, 11-9, and AC 27 MG 14-1 are added in Chapter 3. 4. WEBSITE AVAILABILITY. To access this AC electronically, log on to http://www.airweb.faa.gov/rgl and then click on AC’s. ### Page 5 AC 27-1B, Chg 1 PAGE CONTROL CHART Remove pages Dated Insert Pages Dated iii thru xix ...........................................9/30/99 A-3 thru A-4 .......................................9/30/99 C-43 thru C-50....................................9/30/99 D-7 thru D-105 ...................................9/30/99 E-9 thru E-10 ......................................9/30/99 E-31 thru E-32 ....................................9/30/99 E-39 thru E-44 ....................................9/30/99 F-1 thru F-91.......................................9/30/99 G-3 thru G-47 .....................................9/30/99 MG 5-3 thru MG 5-6 ..........................9/30/99 MG 6-5 thru MG 6-16 ........................9/30/99 MG 8-5 thru MG 8-6 ..........................9/30/99 none MG 12-21 thru MG 12-24 ..................9/30/99 MG 12-37 thru MG 12-38 ..................9/30/99 none none none none none iii thru xxi ...........................................2/12/03 A-3 thru A-4 .......................................2/12/03 C-43 thru C-86....................................2/12/03 D-7 thru D-112 ...................................2/12/03 E-9 thru E-10......................................2/12/03 E-31 thru E-32 ....................................2/12/03 E-39 thru E-44 ....................................2/12/03 F-1 thru F-102 ....................................2/12/03 G-3 thru G-48 .....................................2/12/03 MG 5-3 thru MG 5-6 ..........................2/12/03 MG 6-5 thru MG 6-16 ........................2/12/03 MG 8-5 thru MG 8-6 ..........................2/12/03 MG 11-1 thru MG 11-28 ....................2/12/03 MG 12-21 thru MG 12-24 ..................2/12/03 MG 12-37 thru MG 12-38 ..................2/12/03 MG 13-1 thru MG 13-8 ......................2/12/03 MG 14-1 thru MG 14-8 ......................2/12/03 MG 15-1 thru MG 15-18 ....................2/12/03 MG 16-1 thru MG 16-16 ....................2/12/03 Apdx A-1 thru Apdx A-70 .................2/12/03 David A. Downey Manager, Rotorcraft Directorate Aircraft Certification Service 2 ### Page 6 U.S. Department of Transportation Federal Aviation Administration Subject: CERTIFICATION OF NORMAL CATEGORY ROTORCRAFT Date: Initiated by: 9/30/99 ASW-110 AC No: Change: 27-1B ______________________________________________ 1. PURPOSE: a. This is a total revision of AC 27-1A dated 7/30/97, with Change 1 dated 9/30/98, incorporated. In addition, new material plus changes to existing paragraphs are incorporated. This consolidated version is now renumbered as AC 27-1B and replaces AC 27-1A in its entirety. This revises existing material in 10 paragraphs, adds new material for three paragraphs, and renumbers paragraphs to correspond with Federal Aviation Regulations (FAR) numbering. b. Requests from the rotorcraft industry to make the document easier to use resulted in renumbering the AC paragraphs to correspond with FAR numbering. The figure numbers are also renumbered accordingly. c. This AC does not change regulatory requirements and does not authorize changes in, or deviations from, regulatory requirements. This AC establishes an acceptable means, but not the only means, of compliance. Since the guidance material presented in this AC is not regulatory, terms having a mandatory definition, such as “shall” and “must,” etc., as used in this AC, apply either to the reiteration of a regulation itself, or to an applicant who chooses to follow a prescribed method of compliance without deviation. d. This advisory circular provides information on methods of compliance with 14 CFR Part 27, which contains the Airworthiness Standards for Normal Category Rotorcraft. It includes methods of compliance in the areas of basic design, ground tests, and flight tests. 2. CANCELLATION. AC 27-1A, Certification of Normal Category Rotorcraft, dated 7/30/97, is canceled in its entirety. 3. BACKGROUND. Based largely on precedents set during rotorcraft certification programs spanning over 40 years, this AC consolidates guidance contained in earlier correspondence among FAA headquarters, foreign authorities, the rotorcraft industry, and certificating regions. 4. PRINCIPAL CHANGES: a. Chapter 3 is now titled "Miscellaneous Guidance (MG), Normal Category Rotorcraft," with the following changes: • Paragraphs that correspond to a FAR number are merged into existing AC text in Chapter 2. • Paragraphs that do not correspond with a FAR number either remain in Chapter 3 and are renumbered as MG paragraphs, or are now an appendix. • In order to stay aligned with FAR numbering, Appendices A and C are reserved for future AC material. ### Page 7 b. Paragraphs revised to incorporate technical guidance are AC 21.35, 27.561, 27.952, 27.1305, 27.1309, 27.1351, 27.1353, MG 2, and MG 4 (FADEC). AC 27.661, Rotorblade Clearance, contains new material added as a result of National Transportation Safety Board (NTSB) recommendations. c. New paragraphs added are AC 27.1B and 27.2A, Gross Weight; AC 27.602, Critical Parts; and AC MG 12, External Loads. These paragraphs correspond with recent harmonized regulatory changes. d. The AC is now divided by Subparts and page numbers reflect the relevant FAR Subpart. e. “FAA/AUTHORITY” as used in this document means FAA or another airworthiness authority that has adopted this AC as a means of compliance with the appropriate regulation referenced. 5. DEVIATIONS. As rotorcraft designs vary from conventional configurations, it may become necessary to deviate from the methods and procedures outlined in this AC. These procedures are only one acceptable means of compliance with Part 27. Any alternate means proposed by an applicant will be given due consideration. Applicants are encouraged to use their technical ingenuity and resourcefulness to develop more efficient and less costly methods of achieving the objectives of Part 27. Regulatory personnel and designees should respond to such efforts by the use of engineering judgment in fostering any such efforts as long as the letter and spirit of Part 27 and the Federal Aviation Act are respected. It is recommended that unusual or unique projects be coordinated a sufficient time in advance with the Rotorcraft Standards Staff, ASW-110, or with the appropriate airworthiness authority, to ensure timely and uniform consideration. 6. APPLICABILITY. This material is not to be construed as having any legally binding status and must be treated as advisory only. However, to ensure standardization in the certification process, these procedures should be considered during all rotorcraft type certification and supplemental type certification activities. 7. PARAGRAPHS KEYED TO FAR PART 27. Each paragraph has the applicable amendment to Part 27 shown in the title. All of the original guidance material has been retained as appropriate, even as changes are made to the regulations. This is accomplished through the use of “A,” “B,” etc., paragraphs that follow the original numbered paragraphs. These subsequent paragraphs provide updated guidance information or changes to policy that parallel a specific rule change. The guidance material in the original paragraph (for earlier amendments) still applies and is modified as explained in each of the later paragraphs for later amendments. The applicable amendment number will only appear in the title line for the “A,” “B,” etc., paragraphs. The guidance material in the initial paragraph is intended to apply to all amendments except as modified by the later paragraphs. Each ensuing “A,” “B,” etc., paragraph will be identified with an amendment level to indicate the rule change that precipitated the policy change. 8. RELATED PUBLICATIONS. FAA Certification personnel and designees should be familiar with Order 8110.4, Type Certification, and Order 8100.5, Aircraft Certification Directorate Procedures. Eric Bries Acting Manager, Rotorcraft Directorate Aircraft Certification Service ### Page 8 **CONTENTS** **CHAPTER 1 - FAR 21 CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS** Page No. AC 21.16 Special Conditions ....................................................................................... FAR 21-1 AC 21.31 Type Design................................................................................................. FAR 21-2 AC 21.33 Inspection and Tests.................................................................................... FAR 21-2 AC 21.35 Flight Tests .................................................................................................. FAR 21-2 AC 21.39 Flight Test Instrument Calibration ................................................................ FAR 21-6 and Correction Report CHAPTER 2 - FAR 27 AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT **SUBPART A - GENERAL** **GENERAL** AC 27.1 Applicability … A-1 AC 27.1A Applicability … A-1 AC 27.1B Applicability … A-2 AC 27.2 Special Retroactive Requirements … A-2 AC 27.2A Special Retroactive Requirements … A-4 **SUBPART B - FLIGHT** **GENERAL** AC 27.21 Proof of Compliance … B-1 AC 27.25 Weight Limits … B-3 AC 27.27 Center of Gravity Limits … B-4 AC 27.29 Empty Weight and Corresponding Center of Gravity … B-6 AC 27.31 Removable Ballast … B-7 AC 27.33 Main Rotor Speed and Pitch Limits … B-7 **PERFORMANCE** AC 27.45 Performance - General … B-9 AC 27.45A General … B-18 AC 27.45B General … B-18 AC 27.51 Takeoff … B-25 ### Page 9 SUBPART B - FLIGHT (continued) PERFORMANCE (continued) Page No. AC 27.65 Climb: All Engines Operating … B-26 AC 27.65A Climb: All Engines Operating … B-28 AC 27.67 Climb: One Engine Inoperative … B-28 AC 27.71 Glide Performance … B-29 AC 27.73 Performance at Minimum Operating Speed … B-32 AC 27.75 Landing … B-34 AC 27.79 Limiting Height - Speed Envelope … B-37 AC 27.79A Limiting Height - Speed Envelope … B-45 **FLIGHT CHARACTERISTICS** AC 27.141 Flight Characteristics - General … B-46 AC 27.143 Controllability and Maneuverability … B-47 AC 27.151 Flight Controls … B-52 AC 27.161 Trim Control … B-53 AC 27.171 Stability: General … B-53 AC 27.173 Static Longitudinal Stability … B-54 AC 27.175 Demonstration of Static Longitudinal Stability … B-55 AC 27.177 Static Directional Stability … B-59 **GROUND AND WATER HANDLING CHARACTERISTICS** AC 27.231 Ground and Water Handling Characteristics - General … B-61 AC 27.235 Taxiing Condition..................................................................................................B- 61 AC 27.239 Spray Characteristics … B-62 AC 27.241 Ground Resonance … B-63 **MISCELLANEOUS FLIGHT REQUIREMENTS** AC 27.251 Vibration … B-68 **SUBPART C - STRENGTH REQUIREMENTS** **GENERAL** AC 27.301 Loads … C-1 AC 27.303 Factor of Safety … C-1 AC 27.305 Strength and Deformation … C-2 AC 27.307 Proof of Structure … C-2 ### Page 10 SUBPART C - STRENGTH REQUIREMENTS (continued) GENERAL (continued) Page No. AC 27.307A Proof of Structure.................................................................................................. .C - 3 AC 27.309 Design Limitations … C-4 **FLIGHT LOADS** AC 27.321 Flight Loads - General … C-5 AC 27.337 Limit Maneuvering Load Factor … C-5 AC 27.339 Resultant Limit Maneuvering Loads … C-8 AC 27.341 Gust Loads … C-8 AC 27.351 Yawing Conditions … C-9 AC 27.361 Engine Torque … C-14 **CONTROL SURFACE AND SYSTEM LOADS** AC 27.391 Control Surface and System Loads - General … C-15 AC 27.391A Control Surface and System Loads - General … C-15 AC 27.395 Control System … C-15 AC 27.395A Control System … C-17 AC 27.397 Limit Pilot Forces and Torques … C-17 AC 27.399 Dual Control System … C-18 AC 27.401 Auxiliary Rotor Assemblies … C-18 AC 27.401A Auxiliary Rotor Assemblies … C-20 AC 27.403 Auxiliary Rotor Attachment Structure … C-20 AC 27.403A Auxiliary Rotor Attachment Structure … C-21 AC 27.411 Ground Clearance: Tail Rotor Guard … C-21 AC 27.413 Stabilizing and Control Surfaces … C-22 AC 27.413A Stabilizing and Control Surfaces … C-24 AC 27.427 Unsymmetrical Loads … C-24 **GROUND LOADS** AC 27.471 Ground Loads - General … C-28 AC 27.473 Ground Loading Conditions and Assumptions … C-28 AC 27.475 Tires and Shock Absorbers … C-29 AC 27.477 Landing Gear Arrangement … C-29 AC 27.479 Level Landing Conditions … C-29 AC 27.481 Tail-Down Landing Conditions … C-30 AC 27.483 One-Wheel Landing Conditions … C-31 AC 27.485 Lateral Drift Landing Conditions … C-31 AC 27.493 Braked Roll Conditions … C-32 ### Page 11 AC 27-1B, Chg 2 4/25/06 SUBPART C - STRENGTH REQUIREMENTS( continued) GROUND LOADS (continued) Page No. AC 27.497 Ground Loading Conditions: … C-32 Landing Gear With Tail Wheels AC 27.501 Ground Loading Conditions: … C-34 Landing Gear With Skids AC 27.501A Ground Loading Conditions: … C-35 Landing Gear With Skids AC 27.505 Ski Landing Conditions … C-35 **WATER LOADS** AC 27.521 Float Landing Conditions … C-37 **MAIN COMPONENT REQUIREMENTS** AC 27.547 Main Rotor Structure … C-38 AC 27.549 Fuselage, Landing Gear, and Rotor Pylon Structure … C-38 **EMERGENCY LANDING CONDITIONS** AC 27.561 Emergency Landing Conditions - General … C-40 AC 27.561A Emergency Landing Conditions - General … C-41 AC 27.561B Emergency Landing Conditions - General … C-42 AC 27.561C Emergency Landing Conditions - General … C-43 AC 27.562 Emergency Landing Dynamic Conditions … C-43 AC 27.563 Structural Ditching Provisions … C-81 AC 27.563A Structural Ditching Provisions … C-81 **FATIGUE EVALUATION** AC 27.571 Fatigue Evaluation of Flight Structure … C-83 AC 27.571A Fatigue Evaluation of Flight Structure for Category A Certification … C-86 **SUBPART D - DESIGN AND CONSTRUCTION** **GENERAL** Page No. AC 27.601 Design … D-1 AC 27.602 Critical Parts … D-2 AC 27.603 Materials … D-4 AC 27.605 Fabrication Methods … D-4 AC 27.607 Fasteners … D-5 ### Page 12 **SUBPART D - DESIGN AND CONSTRUCTION** GENERAL (continued) Page No. AC 27.609 Protection of Structure … D-5 AC 27.610 Lightning Protection … D-6 AC 27.610A Lightning and Static Electricity Protection … D-9 AC 27.611 Inspection Provisions … D-13 AC 27.613 Material Strength Properties … D-14 and Design Values AC 27.613A Material Strength Properties … D-16 and Design Values AC 27.619 Special Factors … D-16 AC 27.621 Casting Factors … D-17 AC 27.623 Bearing Factors … D-27 AC 27.625 Fitting Factors … D-28 AC 27.625A Fitting Factors … D-29 AC 27.629 Flutter … D-29 **ROTORS** AC 27.653 Pressure Venting and Drainage of Rotor Blades … D-31 AC 27.659 Mass Balance … D-31 AC 27.661 Rotor Blade Clearance … D-31 AC 27.663 Ground Resonance Prevention Means … D-33 AC 27.663A Ground Resonance Prevention Means … D-35 **CONTROL SYSTEMS** AC 27.671 Control Systems - General … D-36 AC 27-672 Stability Augmentation, Automatic, … D-36 and Power-Operated Systems AC 27.673 Primary Flight Control … D-37 AC 27.674 Interconnected Controls … D-37 AC 27.675 Stops … D-38 AC 27.679 Control System Locks … D-39 AC 27.681 Limit Load Static Tests … D-40 AC 27.683 Operation Tests … D-41 AC 27.685 Control System Details … D-41 AC 27.685A Control System Details … D-42 AC 27.687 Spring Devices … D-43 AC 27.691 Autorotation Control Mechanism … D-43 AC 27.695 Power Boost and Power-Operated Control System … D-44 ### Page 13 AC 27-1B, Chg 2 4/25/06 SUBPART D - DESIGN AND CONSTRUCTION (continued) Page No. **LANDING GEAR** AC 27.723 Shock Absorption Tests … D-47 AC 27.725 Limit Drop Test … D-50 AC 27.727 Reserve Energy Absorption Drop Test … D-50 AC 27.727A Reserve Energy Absorption Drop Test … D-51 AC 27.729 Retracting Mechanism … D-51 AC 27.731 Wheels … D-54 AC 27.733 Tires … D-55 AC 27.735 Brakes … D-56 AC 27.737 Skis … D-58 **FLOATS AND HULLS** AC 27.751 Main Float Buoyancy … D-59 AC 27.753 Main Float Design … D-60 AC 27.755 Hulls … D-60 **PERSONNEL AND CARGO ACCOMMODATIONS** AC 27.771 Pilot Compartment … D-62 AC 27.773 Pilot Compartment View … D-63 AC 27.775 Windshields and Windows … D-68 AC 27.775A Windshields and Windows … D-68 AC 27.777 Cockpit Controls … D-68 AC 27.779 Motion and Effect of Cockpit Controls … D-69 AC 27.783 Doors … D-72 AC 27.783A Doors … D-74 AC 27.785 Seats, Berths, Safety Belts, and Harnesses … D-74 AC 27.785A Seats, Berths, Safety Belts, and Harnesses … D-78 AC 27.785B Seats, Berths, Litters, Safety Belts, and Harnesses … D-80 AC 27.787 Cargo and Baggage Compartments … D-80 AC 27.787A Cargo and Baggage Compartments … D-82 AC 27.801 Ditching … D-82 AC 27.805 Flight Crew Emergency Exits … D-91 AC 27.807 Emergency Exits … D-91 AC 27.807A Emergency Exits … D-93 AC 27.807B Emergency Exits … D-93 AC 27.831 Ventilation … D-94 AC 27.833 Heaters … D-96 **FIRE PROTECTION** AC 27.853 Compartment Interiors … D-98 AC 27.853A Compartment Interiors … D-99 AC 27.855 Cargo and Baggage Compartments … D-99 SUBPART D - DESIGN AND CONSTRUCTION (continued) ### Page 14 FIRE PROTECTION (continued) Page No. AC 27.859 Heating Systems … D-101 AC 27.861 Fire Protection of Structure, Controls, and Other Parts … D-105 AC 27.861A Fire Protection of Structure, Controls, and Other Parts … D-106 AC 27.863 Flammable Fluid Fire Protection … D-107 **EXTERNAL LOAD ATTACHING MEANS** AC 27.865 External Load Attaching Means … D-109 AC 27.865A External Load Attaching Means … D-110 AC 27.865B External Load Attaching Means (previously MG 12) … D-111 **MISCELLANEOUS (DESIGN AND CONSTRUCTION)** AC 27.871 Leveling Marks … D-133 AC 27.873 Ballast Provisions … D-133 **SUBPART E - POWERPLANT** **GENERAL** AC 27.901 Installation … E-1 AC 27.901A Installation … E-5 AC 27.903 Engines … E-5 AC 27.903A Engines … E-7 AC 27.907 Engine Vibration … E-9 **ROTOR DRIVE SYSTEM** AC 27.917 Design … E-10 AC 27.921 Rotor Brake … E-11 AC 27.923 Rotor Drive System and Control Mechanism Tests … E-12 AC 27.923A Rotor Drive System and Control Mechanism Tests … E-17 AC 27.923B Rotor Drive System and Control Mechanism Tests … E-18 AC 27.927 Additional Tests … E-19 AC 27.927A Additional Tests … E-21 AC 27.931 Shafting Critical Speeds … E-21 AC 27.935 Shafting Joints … E-26 AC 27.939 Turbine Engine Operating Characteristics … E-26 ### Page 15 SUBPART E - POWERPLANT (continued) FUEL SYSTEM (continued) Page No. AC 27.951 Fuel System - General … E-29 AC 27.952 Fuel System Crash Resistance … E-31 AC 27.953 Fuel System Independence … E-56 AC 27.954 Fuel System Lightning Protection … E-57 AC 27.955 Fuel Flow … E-58 AC 27.955A Fuel Flow … E-62 AC 27.959 Unusable Fuel Supply … E-64 AC 27.961 Fuel System Hot Weather Operation … E-65 AC 27.961A Fuel System Hot Weather Operation … E-67 AC 27.963 Fuel Tanks: General … E-67 AC 27.963A Fuel Tanks: General … E-68 AC 27.963B Fuel Tanks: General … E-68 AC 27.965 Fuel Tank Tests....................................................................................................E- 69 AC 27.967 Fuel Tank Installation … E-70 AC 27.969 Fuel Tank Expansion Space … E-72 AC 27.969A Fuel Tank Expansion Space … E-72 AC 27.971 Fuel Tank Sump … E-73 AC 27.971A Fuel Tank Sump … E-74 AC 27.973 Fuel Tank Filler Connection … E-74 AC 27.973A Fuel Tank Filler Connection … E-75 AC 27.975 Fuel Tank Vents … E-76 AC 27.975A Fuel Tank Vents … E-77 AC 27.975B Fuel Tank Vents … E-77 AC 27.977 Fuel Tank Outlet … E-78 **FUEL SYSTEM COMPONENTS** AC 27.991 Fuel Pumps … E-80 AC 27.991A Fuel Pumps … E-81 AC 27.993 Fuel System Lines and Fittings … E-81 AC 27.995 Fuel Valves … E-82 AC 27.997 Fuel Strainer or Filter............................................................................................E- 82 AC 27.997A Fuel Strainer or Filter............................................................................................E -83 AC 27.999 Fuel System Drains … E-84 AC 27.999A Fuel System Drains … E-85 **OIL SYSTEM** AC 27.1011 Oil System - General … E-86 AC 27.1013 Oil Tanks … E-86 AC 27.1015 Oil Tank Tests … E-87 AC 27.1017 Oil Lines and Fittings … E-88 ### Page 16 SUBPART E - POWERPLANT (continued) OIL SYSTEM (continued) Page No. AC 27.1019 Oil Strainer or Filter … E-88 AC 27.1019A Oil Strainer or Filter … E-89 AC 27.1021 Oil System Drains … E-90 AC 27.1027 Transmission and Gearboxes: General … E-90 **COOLING** AC 27.1041 Cooling - General … E-92 AC 27.1041A General … E-94 AC 27.1043 Cooling Tests … E-94 AC 27.1045 Cooling Test Procedures … E-97 AC 27.1045A Cooling Test Procedures … E-98 **INDUCTION SYSTEM** AC 27.1091 Air Induction … E-99 AC 27.1091A Air Induction … E-100 AC 27.1093 Induction System Icing Protection … E-100 AC 27.1093A Induction System Icing Protection … E-111 **EXHAUST SYSTEM** AC 27.1121 General … E-113 AC 27.1123 Exhaust Piping … E-114 **POWERPLANT CONTROLS AND ACCESSORIES** AC 27.1141 Powerplant Controls: General … E-117 AC 27.1141A Powerplant Controls: General … E-118 AC 27.1143 Engine Controls … E-118 AC 27.1143A Engine Controls … E-119 AC 27.1143B Engine Controls … E-119 AC 27.1145 Ignition Switches … E-120 AC 27.1147 Mixture Controls … E-120 AC 27.1151 Rotor Brake Controls … E-121 AC 27.1163 Powerplant Accessories … E-122 **POWERPLANT FIRE PROTECTION** AC 27.1183 Lines, Fittings, and Components … E-124 AC 27.1185 Flammable Fluids … E-125 AC 27.1187 Ventilation … E-126 ### Page 17 AC 27-1B, Chg 2 4/25/06 SUBPART E - POWERPLANT (continued) POWERPLANT FIRE PROTECTION (continued) Page No. AC 27.1189 Shutoff Means … E-127 AC 27.1191 Firewalls … E-128 AC 27.1193 Cowling and Engine Compartment Covering … E-132 AC 27.1193A Cowling and Engine Compartment Covering … E-133 AC 27.1194 Other Surfaces … E-134 AC 27.1195 Fire Detector Systems … E-135 **SUBPART F - EQUIPMENT** **GENERAL** AC 27.1301 Function and Installation.........................................................................................F- 1 AC 27.1303 Flight and Navigation Instruments … F-1 AC 27.1305 Powerplant Instruments … F-7 AC 27.1305A Powerplant Instruments … F-8 AC 27.1305B Powerplant Instruments … F-9 AC 27.1307 Miscellaneous Equipment … F-11 AC 27.1309 Equipment, Systems, and Installations … F-13 **INSTRUMENTS - INSTALLATION** AC 27.1321 Arrangement and Visibility … F-34 AC 27.1322 Warning, Caution, and Advisory Lights … F-34 AC 27.1323 Airspeed Indicating System … F-39 AC 27.1325 Static Pressure Systems … F-40 AC 27.1327 Magnetic Direction Indicator … F-41 AC 27.1329 Automatic Pilot System … F-41 AC 27.1335 Flight Director Systems … F-50 AC 27.1337 Powerplant Instruments … F-51 AC 27.1337A Powerplant Instruments … F-52 **ELECTRICAL SYSTEMS AND EQUIPMENT** AC 27.1351 Electrical Systems and Equipment - General … F-54 AC 27.1353 Storage Battery Design and Installation … F-58 AC 27.1357 Circuit Protective Devices … F-62 AC 27.1361 Master Switch … F-64 AC 27.1365 Electric Cables … F-64 AC 27.1367 Switches … F-66 ### Page 18 SUBPART F - EQUIPMENT (continued) **LIGHTS** Page No. AC 27.1381 Instrument Lights … F-67 AC 27.1383 Landing Lights … F-67 AC 27.1385 Position Light System Installation … F-68 AC 27.1387 Position Light System Dihedral Angles … F-68 AC 27.1389 Position Light Distribution and Intensities … F-68 AC 27.1391 Minimum Intensities in the Horizontal … F-68 Plane of Forward and Rear Position Lights AC 27.1393 Minimum Intensities in Any Vertical … F-68 Plane of Forward and Rear Position Lights AC 27.1395 Maximum Intensities In Overlapping … F-68 Beams of Forward and Rear Position Lights AC 27.1397 Color Specifications … F-68 AC 27.1399 Riding Light … F-68 AC 27.1401 Anti-collision Light System … F-70 **SAFETY EQUIPMENT** AC 27.1411 General … F-71 AC 27.1413 Safety Belts … F-73 AC 27.1415 Ditching Equipment … F-74 AC 27.1419 Ice Protection … F-75 **MISCELLANEOUS EQUIPMENT** AC 27.1435 Hydraulic Systems … F-97 AC 27.1457 Cockpit Voice Recorder … F-99 AC 27.1459 Flight Recorders … F-101 AC 27.1461 Equipment Containing High Energy Rotors … F-103 **SUBPART G - OPERATING LIMITATIONS AND INFORMATION** **OPERATING LIMITATIONS** AC 27.1501 Operating Limitations - General … G-1 AC 27.1503 Airspeed Limitations: General … G-1 AC 27.1505 Never-Exceed Speed … G-4 AC 27.1509 Rotor Speed … G-6 AC 27.1519 Weight and Center of Gravity … G-8 AC 27.1521 Powerplant Limitations … G-9 AC 27.1521A Powerplant Limitations … G-10 AC 27.1521B Powerplant Limitations … G-10 ### Page 19 AC 27-1B, Chg 2 4/25/06 SUBPART G - OPERATING LIMITATIONS AND INFORMATION (continued) OPERATING LIMITATIONS (continued) Page No. AC 27.1523 Minimum Flightcrew … G-11 AC 27.1525 Kinds of Operations … G-15 AC 27.1527 Maximum Operating Altitude … G-16 AC 27.1529 Instructions for Continued Airworthiness … G-16 (Maintenance Manual) AC 27.1529A Instructions for Continued Airworthiness … G-20 (Maintenance Manual) **MARKINGS AND PLACARDS** AC 27.1541 Marking and Placards - General … G-21 AC 27.1543 Instrument Markings: General … G-21 AC 27.1545 Airspeed Indicator … G-25 AC 27.1547 Magnetic Direction Indicator … G-25 AC 27.1549 Powerplant Markings … G-25 AC 27.1549A Powerplant Markings … G-26 AC 27.1551 Oil Quantity Indicators … G-26 AC 27.1553 Fuel Quantity Indicator … G-26 AC 27.1555 Control Markings … G-27 AC 27.1557 Miscellaneous Markings and Placards … G-28 AC 27.1559 Limitations Placard … G-30 AC 27.1561 Safety Equipment … G-31 AC 27.1565 Tail Rotor … G-32 **ROTORCRAFT FLIGHT MANUAL** AC 27.1581 Rotorcraft Flight Manual - General … G-33 AC 27.1583 Operating Limitations … G-40 AC 27.1585 Operating Procedures … G-42 AC 27.1585A Operating Procedures … G-45 AC 27.1587 Performance Information … G-45 AC 27.1589 Loading Information … G-48 ### Page 20 **CHAPTER 3 - MISCELLANEOUS GUIDANCE (MG),** **NORMAL CATEGORY ROTORCRAFT** AC 27 MG 1 Certification Procedure for Rotorcraft Avionics Equipment..............................MG 1-1 AC 27 MG 2 Standardized Test Procedure for Rotorcraft DC ..............................................MG 2-1 Electrical System Tests AC 27 MG 3 Definition of Engine Isolation Features as Applied to ......................................MG 3-1 §§ 27.79(b)(2), 27.141(b)(1), and 27.143(d)(1) AC 27 MG 4 Full Authority Digital Electronic Controls (FADEC) ..........................................MG 4-1 AC 27 MG 5 Agricultural Dispensing Equipment Installation................................................MG 5-1 AC 27 MG 6 Emergency Medical Service (EMS) Systems, Installations,.............................MG 6-1 Interior Arrangements, and Equipment AC 27 MG 7 Hour Meters or Time-in-Service Recording Devices in Rotorcraft ...................MG 7-1 AC 27 MG 8 Substantiation of Composite Rotorcraft Structure............................................MG 8-1 AC 27 MG 9 Rotorcraft One-Engine-Inoperative (OEI) Power Assurance ...........................MG 9-1 AC 27 MG 10 Advisory Material for Substantiation of Emergency .......................................MG 10-1 Flotation System AC 27 MG 11 Fatigue Evaluation of Rotorcraft Structure.....................................................MG 11-1 AC 27 MG 12 External Loads (see AC 27.865B) .................................................................MG 12-1 AC 27 MG 13 Systems Certification Considerations ............................................................MG 13-1 AC 27 MG 14 Certification Procedure for Installation of Vapor Cycle Air Conditioning ........MG 14-1 Systems AC 27 MG 15 Airworthiness Approval of Rotorcraft Health Usage Monitoring.....................MG 15-1 Systems (HUMS) AC 27 MG 16 Certification Procedure for Rotorcraft Night Vision Imaging ..........................MG 16-1 Systems (NVIS) Equipment AC 27 MG 17 (reserved) ......................................................................................................MG 17-1 AC 27 MG 18 Helicopter Terrain Awareness and Warning System (HTAWS) ....................MG 18-1 AC 27 Appendix A - Instructions for Continued Airworthiness............................................... Apdx A-1 AC 27 Appendix B - Airworthiness Guidance for Rotorcraft Instrument Flight....................... Apdx B-1 AC 27 Appendix C - (reserved) (Criteria for Category A) .....................................................Apdx C-1 ### Page 21 AC 27-1B, Chg 2 4/25/06 **FIGURES** The first digits in the figure number are the same as the related paragraph number. The digits after the hyphen are the number within the paragraph. Most figures are located at the end of the related paragraph. Figure No. Title Page AC 27.45-1 H-V Requirements B-19 AC 27.45-2 IGE Controllability Requirements B-20 AC 27.45-3 Shaft Horsepower vs. Turbine Outlet B-21 Temperature - Sea Level Standard Day AC 27.45-4 Uninstalled Takeoff Power Available B-22 AC 27.45-5 Installed Takeoff Power Available, Anti-Ice Off, B-23 400 RPM AC 27.45-6 Power Assurance Check Chart B-24 AC 27.71-1 Autorotational Characteristics - Typical B-31 AC 27.79-1 Height-Velocity Diagram B-43 AC 27.79-2 Altitude/Wei ght Accountability B-44 AC 27.175-1 Static Longi tudinal Stability B-58 AC 27.251-1 Demonstration Points B-70 AC 27.337-1 Load Factor Gross Weight Curve C-7 AC 27.351-1 Sample Yaw/Forward Speed Diagram C-12 AC 27.351-2 Safety Factors for Probability of Failure C-13 AC 27.427-1 View Looking Forward C-26 AC 27.427-2 View Looking Forward C-27 AC 27.562-1 Seat Restraint System Dynamic Tests C-47 AC 27.562-2 Schematic Floor Defo rmation Fixture; Seat Legs C-54 Attached at Floor Level ### Page 22 Figure No. Title Page AC 27.562-3 Schematic Test Fixt ure; Side Wall Mounted Seat C-58 AC 27.562-4 Installation of Pe lvic-Lumbar Spine Load Cell C-62 in Part 572B Anthropomorphic Dummy AC 27.562-5 Measurement to be Ta ken Over Full Width of C-74 Seat Bottom Cushion AC 27.562-6 Post Test Condition C-74 AC 27.562-7 Impulse Shape C-80 AC 27.562-8 Change in Velocity C-80 AC 27.571-1 Comparison Percent of Time in Level Flight C-85 AC 27.621-1 Example of Cast ing Load Sheet Retract D-24 Actuator Support-Landing Gear AC 27.621-2 Example of Casti ngs to be Static Tested D-26 for Model XYZ AC 27.773-1 Cockpit Visibility D-67 AC 27.801-1 Sea State Code - World Meteorological D-90 Organization AC 27.865B-1 U.S. Operational (Part 133) Rotorcraft-Load D-116 Combination Versus Applicable Cargo Type Data and Definition Summary AC 27.931-1 Cantilevered Shaft, First Critical Speed E-25 AC 27.931-2 Shaft Between Support Bearings, First Critical E-25 Speed AC 27.952-1 Static Tensi on and Shear Loading Modes E-50 AC 27.952-2 Hose Assembly Tests E-51 AC 27.952-3 Minimum Average and Individual Loads for Hose E-52 and Hose-End Fitting Combinations AC 27.952-4 Typical Method of Breakaway Fuel Fitting Load E-53 Calculations ### Page 23 AC 27-1B, Chg 2 4/25/06 Figure No. Title Page AC 27.952-5 Typical Methods of Frangible or Deformable E-54 Attachment Load Calculations, Frangible Bolts AC 27.952-6 Typical Methods of Frangible or Deformable E-55 Attachment Load Calculations, Frangible Baffle AC 27.955-1 Fuel Flow E-61 AC 27.1041-1 Additional Cooling Test Point E-93 AC 27.1191-1 Table of Materials and Gauges Acceptable for E-131 Fireproof Protective Devices with Flat Surface Geometries AC 27.1329-1 Deviation Profile F-48 AC 27.1329-2 Operational Limitation F-49 AC 27.1419-1 Continuous Icing-Temper ature vs. Altitude Limits F-93 AC 27.1419-2 Intermittent Icing-Tem perature vs. Altitude Limits F-94 AC 27.1419-3 Continuous Icing-Li quid Water Content vs Drop F-95 Diameter AC 27.1419-4 Intermittent Icing-Li quid Water Content vs. Drop F-96 Diameter AC 27 MG 3-1 Part 29 Engine Isolation Rules MG 3-5 AC 27 MG 5-1 Acceptable Ultim ate Load Factor for MG 5-2 Agricultural Dispensing Equipment Design AC 27 MG 5-2 Sketch of Tank Pressure Test MG 5-6 AC 27 MG 6-1 Typical Liquid Oxygen System MG 6-17 AC 27 MG 6-2 Typical Liquid Oxygen System - Using MG 6-18 Combination Valve AC 27 MG 8-1 Schematic Diagram of Building Block Tests MG 8-16 AC 27 MG 8-2 Characterization of Impact Damage MG 8-20 ### Page 24 Figure No. Title Page AC 27 MG 8-3 Residual Strength vs. Time MG 8-23 AC 27- MG 8-4 Residual Str ength and Damage Size MG 8-24 Relationships for Fail-Safe Substantiation AC 27 MG 9-1 Installed Engine Fuel Flow Capacity Check MG 9-8 AC 27 MG 11-1 Example of Variation in Time Spent in Level MG 11-7 Flight for Three Rotorcraft Types AC 27 MG 11-2 Flight Regime to be Investigated Power-On MG 11-17 Operation AC 27 MG 11-3 Flight Regime to be Investigated for Power-Off MG 11-17 Operation AC 27 MG 11-4 Simplified Analytical Method for Safe Life MG 11-18 Evaluation AC 27 MG 11-5 Typical S/N Curve for Safe Life Evaluation MG 11-18 AC 27 MG 11-6 Elements of Safe Life Determination MG 11-19 AC 27 MG 11-7 Flight Load Measurem ent Program Conditions MG 11-20 to be Investigated AC 27 MG 11-8 Lifting Helicopt er Load Measurement MG 11-23 AC 27 MG 11-9 Example of Twin Turbine Usage Spectrum MG 11-24 AC 27 MG 14-1 Refrigeration Cycle MG 14-7 AC 27 MG 15-1 Flow Chart for Use of Mitigating Actions MG 15-8 Applied for Reduction of Software Level AC 27 MG 15-2 Flow Chart for App lication of HUMS Specific MG 15-11 Software with a Ground Base that uses COTS Software for Operational Software AC 27 MG 18-1 Recommended Alert Prioritization between MG 18-16 HTAWS and TCAS TA AC 27 MG 18-2 Flight Test Matrix MG 18-19 ### Page 25 AC 27-1B, Chg 2 4/25/06 Figure No. Title Page AC 27 MG 18-2.1 TAWS Required Terrain Clearance (RTC) MG 18-25 by Phase of Flight AC 27 APX B-1 Rotorcraft Dynamic Stability Requirements Apdx B-21 for IFR ### Page 26 Abbreviations and Definitions AC advisory circular ACO aircraft certification office ADF automatic direction finding ADI attitude direction indicator AEO all engines operating AFCS automatic flight control systems AGL above ground level AHRS attitude heading reference system Amdt. Amendment APU auxiliary power unit ATC air traffic control BIM blade inspection method CAM cockpit area microphone CAR Civil Air Regulations CAS calibrated air speed CBIM cockpit blade inspection method CDP critical decision point CG center of gravity COTS commercial off-the-shelf CPS cycles per second CRFS crash resistant fuel system CRT cathode ray tube CVR cockpit voice recorder DER designated engineering representative DME distance measuring equipment ECAS engine caution advisory systems ECU environmental control unit EFIS electronic flight instrument system EFP engine failure point EHE exhaust heat exchanger ELT emergency locator transmitter EMC electromagnetic compatibility EMI electromagnetic interference EMS emergency medical service EOL end of life FAA Federal Aviation Administration FADEC full authority digital engine control FAR Federal Aviation Regulations FEM finite element modeling FHA functional hazard assessment FMEA failure mode and effects analysis FPM feet per minute FPS feet per second GPS global positioning system HF high frequency HIRF high intensity radiated fields HRD high rate of discharge HUMS health and usage monitoring systems HV height-velocity IAS indicated airspeed ICA instructions for continued airworthiness ICAO International Civil Aviation Organization ICS inter-communication system IFR instrument flight rules IGE in ground effect IIDS integrated instrument display system ILS instrument landing system IMC instrument meteorological conditions INS inertial navigation system ISA international standard atmosphere ISIS integral spar inspection system ITT inter-turbine temperature KCAS knots calibrated airspeed KIAS knots indicated air speed KTAS knots true airspeed LDP landing decision point LWC liquid water content MCP maximum continuous power MEL minimum equipment list MGT measured gas temperature MMEL master minimum equipment list MSL mean sea level MVD median volume diameter NDI non-destructive inspection NM nautical mile NPRM notice of proposed rulemaking NTSB National Transportation Safety Board NVG night vision goggles Page xxi ### Page 27 AC 27-1B, Chg 1 2/12/03 NVIS night vision imaging systems OAT outside air temperature OEI one engine inoperative OGE out of ground effect PBA pitch bias actuator PCF post crash fire PIO pilot induced oscillation PSA preliminary safety assessment PSIG pounds per square inch gauge QPL qualified parts list RFM rotorcraft flight manual RFMS rotorcraft flight manual supplement RPM revolutions per minute RTCA Radio Technical Commission of Aeronautics RVR runway visual range SAE Society of Automotive Engineers SAS stability augmentation system SCAS stability and control augmentation systems S/N stress vs. number of cycles SRM structural repair manual SSA system safety assessment STC supplemental type certificate STOL short takeoff and landing TBO time between overhaul TC type certificate TCDS type certificate data sheet TDP takeoff decision point TIA type inspection authorization TIR type inspection report TOT turbine outlet temperature TSO: technical standard order TVP true vapor pressure VBIM visual blade inspection method VFR visual flight rules VMC visual meteorological conditions VOR very high frequency omnidirectional range radio VSI vertical speed indicator V/STOL vertical/short takeoff and landing VTOL vertical takeoff and landing WAT weight, altitude, temperature Altitudes HD density altitude HP pressure altitude V speeds VD diving speed VH speed in level flight with maximum continuous power VMO maximum operating limit speed VNE never-exceed speed VTOSS takeoff safety speed for Category A rotorcraft VX speed for best angle of climb VY speed for best rate of climb MMO maximum operating mach number N speeds NF free turbine speed NG gas generator speed NP power turbine speed NR rotor speed Coefficients CD coefficient of drag CL coefficient of lift CP coefficient of power CT coefficient of thrust Page xxii ### Page 28 **CHAPTER 1. PART 21** **FAR 21 - GENERAL** CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS (Amendment 21-50) AC 21.16. § 21.16 SPECIAL CONDITIONS . a. The Process . Chapter 2, Section 1, Paragraph 8 of the Type Certificate Handbook, Order 8110.4A, provides detailed guidance on the special conditions process. However, much of that material has been outdated with the implementation of the Aircraft Certification Directorate Program. Rotorcraft special conditions are processed through the Rotorcraft Standards Staff, ASW-110. That office will ensure coordination with the affected agency and industry elements including the Assistant Chief Counsel. All comments will be considered and the disposition documented by the Rotorcraft Directorate. ASW-100 will issue the special conditions. b. Basis for Development . (1) Special conditions are justified on the basis of the existing Part 27 being inadequate or inappropriate due to novel or unusual design features of the rotorcraft to be certificated. (2) The phrase “novel or unusual” as used in § 21.16 is a very relative term. As used hereafter in applying § 21.16 to justify the issuance of special conditions, “novel or unusual” will be taken with respect to the state of technology envisaged by the applicable airworthiness standards of this subchapter. It must be recognized that in some areas which will vary from time to time, the state of the regulations may somewhat lag the state of the art in new design because of the rapidity in which the state of the art is advancing in civil aeronautical design and because of the time required to develop the experience base needed by the FAA/AUTHORITY to proceed with general rulemaking. Applicants for type certification of a new design have the opportunity to mitigate the impact of not knowing the precise airworthiness standards to be applied for “novel or unusual design features” by consulting with the FAA/AUTHORITY early in their certification planning when such features are suspected or known by the applicant to exist. It should also be recognized that, because of the intentional objective nature of the airworthiness standards of this subchapter, many new design features which might be thought of as “novel or unusual” may already be adequately covered by existing regulations, thus obviating the need to issue special conditions. (3) Before proposing special conditions , the certification staff should very thoroughly analyze the existing regulations and ensure they are inadequate or inappropriate in light of a new and novel design feature. Page FAR 21 - 1 ### Page 29 AC 21.31. § 21.31 TYPE DESIGN . The regulatory basis for requiring data to def ine the design is contained in § 21.31. This section is self-explanatory and broad enough in scope to give the certification staff access to sufficient data to determine compliance with Part 27. AC 21.33. § 21.33 INSPECTION AND TESTS . a. Applicant Responsibility . Section 21.33 requires the applicant to: (1) Ensure the test rotorcraft confo rms to the type design. This must be accomplished prior to presentation to the FAA/AUTHORITY for testing. (2) Conduct all inspections and tests necessary to determine compliance with the airworthiness and noise requirements. b. FAA/AUTHORITY Responsibility . (1) The design evaluation engineers s hould ensure that the type design is adequate in their technical area and that the inspections and tests to be conducted are appropriate and sufficient to show compliance with Part 27. (2) As changes to the rotorcraft are made during the test program, the flight test crew should ensure that the appropriate design evaluation engineer concurs with the change and the conformity inspection of the change has been conducted. AC 21.35 § 21.35 (Amendm ent 21-59) FLIGHT TESTS. a. Explanation . (1) This section outlines the require ments of the applicant for aircraft type certification and should be used in conjunction with FAA Order 8110.4A, Section 5. Section 21.35 requires, in part, that the applicant conduct sufficient flight tests to show compliance with the flight requirements throughout the proposed flight envelope. The results of the applicant’s flight test should be submitted to the FAA/AUTHORITY in report form for evaluation to determine what verification flight tests the FAA/AUTHORITY may elect to conduct. The report should conclude that in the applicant’s opinion the test aircraft complies with the applicable certification requirements. The FAA/AUTHORITY verification flight test should include, but not be limited to, the critical or marginal results contained in the applicant’s flight test report. The FAA/AUTHORITY’s role in the certification effort is not envisioned to be one of conducting day-to-day routine flight tests with the applicant, but only to verify his results through limited sampling. In certain tests, such as high altitude testing at a remote mountain site, there is an advantage in conducting flight tests concurrently with the applicant. Additionally, the FAA/AUTHORITY can provide technical flight test Page FAR 21 - 2 ### Page 30 assistance to the applicant in certain cases. This can be done after a cursory review and a letter of authorization is issued to the flight test crew. (2) Preflight Test Planning . After the applicant’s flight test report is reviewed, it should be determined what FAA/AUTHORITY engineering flight tests are necessary. These tests are normally specified in the Type Inspection Authorization (TIA). At the same time the FAA/AUTHORITY must know and agree to the applicant’s proposed means of data acquisition, reduction, and expansion of the flight test data. The adequacy of the test instrumentation should be evaluated prior to official type certification tests (reference paragraph AC 21.39). (3) Order of Testing . The Federal Aviation Regulations are so worded that the results of some flight tests have a definite bearing on the conduct of other tests. For this reason, and to minimize retesting, careful attention should be given to the order of testing. The exact order of testing will be determined only by considering the particular rotorcraft and test program involved. Tests which are particularly important in the early stages of the program are: (i) Airspeed calibration: All te sts involving airspeed depend upon the calibration. (ii) Engine power available determination. (iii) Engine cooling. (4) Test Groupings . (i) Weight and c.g .: In addition to the regulatory relationship of one test to another, efficient testing requires that consideration be given to the accomplishment of as many tests on a single flight as can be accommodated successfully. (ii) Special Instrumentation . Similarly, consideration should be given to grouping of tests that involve special instrumentation. Examples of these are takeoff and landing tests which usually require group equipment to record horizontal distance, height, and time. Ground calibration of the airspeed indicating system can be accomplished at the same time. It is the applicant’s responsibility to provide the necessary instrumentation. (5) Functional and Reliability Testing (i) Section 21.35(b)(2) requires that the applicant determine that “there is reasonable assurance that the aircraft, its components, and its equipment are reliable and function properly.” Section 21.35(f)(1) requires a Function and Reliability (F&R) program of 300 hours for turbine engine powered aircraft incorporating engines of a type not previously used in a type certificated aircraft. Section 21.35(f)(2) requires a 150-hour F&R program for all other aircraft. The following reflects general practices that Page FAR 21 - 3 ### Page 31 have been used during rotorcraft certification programs. FAA/AUTHORITY have supported proposals which gave F&R test time credit for certification testing in lieu of dedicated F&R testing. In establishing such credit, the following should be considered: (A) The point in time in which the rotorcraft reaches substantial conformity with the approved type design. (B) The extent and comple xity of the new design. (C) For a previously certified rotorcraft, the F&R program requirement should be commensurate with the modification or change in type design and may be zero. (ii) Historically, for major rotorcraft type certification programs, flight time credit has been limited so as to require an irreducible minimum of 50 hours of dedicated F&R flight time. For rotorcraft programs that involved new engine installations (mature engine design) or drive train/rotor system changes on previously certified aircraft (TC amendments or STC’s), flight time credit has been liberal and often resulted in very little or no dedicated F&R testing. b. Procedures . (1) Type Certification Flight Tests . (i) Prior to initiating official F AA/AUTHORITY flight tests, a conformity inspection of the test aircraft must be accomplished. This is needed to assure that the test aircraft is in the proper configuration or “conforms” to the engineering drawings and documents that have been submitted to FAA/AUTHORITY, evaluated, and approved. It is absolutely essential to know the configuration being tested in any engineering flight evaluation. Conformity inspection prior to TIA flight tests assures that testing will not be wasted because of configuration uncertainties. (ii) FAA Order 8110. 4A, paragraph 67, contains a requirement that the applicant must keep the FAA/AUTHORITY advised of any configuration changes to the aircraft. The manufacturing inspector should keep the FAA/AUTHORITY flight test pilot apprised of any change which may affect safety of the test aircraft or may influence test results. (iii) Results of the conformity inspection and the engineering flight test program must be documented. This is normally done in the Type Inspection Report (TIR). Results may be documented in any acceptable engineering format. The report should be in sufficient detail to clearly show how compliance with each appropriate section of the rules was determined. (iv) The flight test pilot mu st assure that the FAA/AUTHORITY manufacturing inspector and certification engineer are aware of all configuration Page FAR 21 - 4 ### Page 32 changes found necessary as a result of FAA/AUTHORITY tests. The manufacturing inspector is responsible for assuring that all changes are incorporated into production drawings after the design data reflecting the change have been approved by the certification engineer. (v) Additional flight test res ponsibilities, procedures, and requirements during the certification flight test process are contained in FAA Order 8110.4A, Section 5, Flight. (2) Function and Reliability Tests . (i) A comprehensive and systematic check of all aircraft components must be made to assure that they perform their intended function and are reliable. (ii) F&R testing should be accompli shed on an aircraft which conforms to the type design. Non-conformities must be documented and accepted. F&R testing should follow the type certification testing described in paragraph AC 21.35b(1) above to assure that significant changes resulting from type certification tests are incorporated on the aircraft prior to F&R tests. (iii) All components of the roto rcraft should be periodically operated in sequences and combinations likely to occur in service. Ground inspections should be made at appropriate intervals to identify potential failure conditions; however, no special maintenance beyond that described in the aircraft maintenance manual should be allowed. (iv) A complete record of defec ts and failures should be maintained along with required servicing of aircraft fluid levels. Results of this record should be consistent with inspection and servicing information provided in the aircraft maintenance manual. (v) A certain portion of the F& R test program may focus on systems, operating conditions, or environments found particularly marginal during type certification tests. (vi) A substantial portion of the flying should be on a si ngle aircraft. The flying should be carried out to an intensive schedule on an aircraft that is very close to the final certification standard, operated and maintained as though it were in service. A range of representative ambient operating conditions and sites should be considered. It is acceptable for non-F&R flight testing conducted at various sites and in varying ambient conditions to be used to satisfy the F&R requirements for those conditions. Page FAR 21 - 5 ### Page 33 AC 21.39. § 21.39 (Amendm ent 21-59) FLIGHT TEST INSTRUMENT CALIBRATION AND CORRECTION REPORT . a. Explanation . It is the applicant’s responsibility to provide instrumentation for all parameters needed to show compliance with the airworthiness regulations. (1) For those data which are necessary to show compliance with the regulations, a permanent record should be established. A permanent record is acceptable in either graphical or photographic form, and in some instances a manual recording may be satisfactory. (2) Regardless of the record form , the accuracy of the record must be established by reference to a laboratory standard traceable to the National Bureau of Standards. (3) If multiplexing is used, the time ba se must be synchronized to a reference point from which the magnitude of each parameter can unquestionably be determined. Also, the sampling rate should be sufficiently frequent to ensure that the maximums, minimums, and trends of magnitude of the parameter are recorded with respect to time. b. Procedures . Prior to conducting flight tests, the FAA/AUTHORITY flight test team should review the applicant’s flight test instrumentation, calibration, and correction report. (1) The frequency of recalibration va ries with the consistency of the instrumentation under consideration. For example, cyclic and collective position is sometimes calibrated immediately before and after a flight where these parameters are used to provide critical flight data. Six months is a typical interval for recording/signal conditioning and nonstrain gage sensors, while 1 year is typical for strain gaged components. Also, environmental effects such as vibration, humidity, temperature, etc., should be considered when determining whether recalibration is necessary. (2) The highest and lowest magnitude of the parameter being recorded should be considered when establishing the scale for instrumentation. Ideally, the highest magnitude throughout the flight would fall on the maximum indicating point of the recording. Page FAR 21 - 6 ### Page 34 **CHAPTER 2. PART 27** **AIRWORTHINESS STANDARDS** **NORMAL CATEGORY ROTORCRAFT** **SUBPART A - GENERAL** AC 27.1. § 27.1 APPLICABILITY. a. Explanation. This section prescribes the rotorcraft categories eligible for certification under this Part. There is no minimum weight limit for certification under Part 29; however, Part 27 is applicable to rotorcraft with maximum weights of 6,000 pounds or less. (1) Without Engine Isolation. For single-engine rotorcraft and multiengine rotorcraft without engine isolation, the height-velocity (HV) diagram is conducted with sudden failure of all engines, and the takeoff maneuver must pass through the clear area of the diagram to the 50-foot point with all engines operating. (2) With Engine Isolation. Part 27 multiengine rotorcraft may be certificated with engine isolation features (reference paragraph AC 27 MG 3). These rotorcraft are not required to meet the Part 29, Category A, performance requirements, and continued flight after an engine failure is not assured since under some conditions failure of the remaining engine may occur after a limited time. The takeoff is conducted with all engines operating, while the height-velocity diagram is determined with the most critical engine inoperative. If complete Part 29, Category A, design features and performance are achieved, the Category A performance may be included in the FAA/AUTHORITY-approved portion of the Rotorcraft Flight Manual although this performance is not required by the regulations. b. Procedures. None. AC 27.1A. § 27.1 (Amendment 27-33) APPLICABILITY. a. Explanation. Amendment 33 formally introduced the requirements for certification of a Part 27 rotorcraft to Category A design and performance standards. These standards are found in Appendix C of Part 27. The establishment of Category A design and performance for multiengine Part 27 rotorcraft is still voluntary. If so requested, the corresponding AC 29-2C material applies. b. Procedures. None Page A - 1 ### Page 35 AC 27.1B § 27.1 (Amendment 27-37) APPLICABILITY. a. Explanation. Amendment 27-37 increases the normal category rotorcraft maximum gross weight limit from 6,000 to 7,000 pounds and introduces new airworthiness requirements. The gross weight increase was intended to compensate for past regulatory actions that, in effect, cumulatively increased the gross weight of normal category rotorcraft. The amendment also adds an explicit nine-passenger seat limitation. The new airworthiness requirements were intended to increase the level of safety to support a potential increase in the number of passenger seats associated with the increase in maximum gross weight limit. b. Procedures. None AC 27.2 § 27.2 (Amendment 27-28) SPECIAL RETROACTIVE REQUIREMENTS. a. Explanation. (1) Amendment 27-28 requires a combined shoulder harness and safety belt (also called a torso restraint system) at each occupant’s seat for all rotorcraft manufactured after September 16, 1992. (2) The design features of the restra int system are mainly contained in this section rather than having to refer to other sections within Part 27 except for a general reference to the differing strength standards between earlier static strength only standards and the static and dynamic strength standards of Amendment 27-25. (3) Combined safety belt and harness strength standards system follows: (i) Those rotorcraft type designs ce rtificated to static strength standards alone prior to Amendment 27-25, such as 4 g’s forward may use belt and harness systems, characterized as 1,500 pounds strength systems, provided they comply with those standards. TSO C22f and earlier restraint systems have such ratings. A combined belt and harness with a 1,500 pounds rating, which comply with the Part 27 standards for the rotorcraft type design, but are not necessarily TSO approved, may be approved as a part of the type design. Such design information for a non-TSO’d item would be included in a note on the aircraft type certificate data sheet (TCDS) or specification sheet by part number as “required equipment.” TSO C114-approved torso restraint systems, characterized as 3,000 pounds strength system, may be used provided the design features comply with this section, but no special information on the TCDS is necessary. (ii) Those rotorcraft type designs ce rtified to dynamic test requirements of Amendment 27-25 should use torso restraint systems approved under TSO C114 or approved under equivalent standards such as those contained in Part 27. Page A - 2 ### Page 36 (4) Load Distribution and Design Requirements. Although not stated in § 27.2, a 60 percent and 40 percent load distribution between the safety belt and harness, respectively, is required in § 27.785(g). The safety belt should withstand 100 percent if the safety belt is capable of being used alone. Also, the safety belt or harness attachments to the seat or structure should include the 1.33 factor described in § 27.785(f)(2) of Amendment 27-21 for those rotorcraft with that certification criteria or should include the 1.15 factor as described in § 27.625 (and predecessor Part 6) standards for those rotorcraft with the earlier certification criteria. A factor is used whether test or analysis methods are used for static substantiation of the seating systems. Refer to paragraph AC 27.785c(i) (§ 27.785). (5) The companion operating rule change of Amendment 91-220, amended § 91.205 (Amendment 91-223), affecting the aircraft equipment requirements. Operating rule § 91.107(a) already requires use of the harness whenever the aircraft seat is so equipped. b. Procedures. (1) A TSO-approved combined safety belt and harness or torso restraint system may be used provided the installation requirements in § 27.2 are satisfied. A combined belt and harness (not necessarily TSO approved) may be approved as a part of the rotorcraft type design and so noted on the aircraft specification or TCDS. (2) Structural analysis or static test may be used. For those rotorcraft designs that are subject to the dynamic test standards of § 27.562, the torso restraint system is required to be qualified for the particular use or installation in each rotorcraft type design. A dynamic test may be required for alternate restraint systems as well as the originally approved system. TSO C114 approval does not constitute approval for installation of a restraint system in a rotorcraft design subject to dynamic test. (i) Paragraph 27.562 of this AC concerns in part the dynamic test standards of Amendment 27-25. (ii) AC 23-4 dated June 20, 1986, c oncerns static test procedures for small airplane seats and restraint systems. (Certain small airplanes manufactured after December 12, 1986, should have harnesses for each seat also.) A test proposal for rotorcraft installations may adopt procedures appropriate to the particular installation. The 60/40 percent distribution is sufficiently achieved when the blocks in Figure 4 of AC 23-4 are used. (iii) The static design side load fo r the harness installation may be proven by test or analysis using the load distribution previously noted. For “older” designs, the side load of § 27.561(b)(3)(iii) is 2.0g, and for later designs (Amendment 27-25 and later), it is 8.0g. Page A - 3 ### Page 37 AC 27.2A § 27.2 (Amendment 27-37) SPECIAL RETROACTIVE REQUIREMENTS. a. Explanation. Amendment 27-37 specifies that for any change in type design to a rotorcraft that increases the passenger-seat capacity to a number greater than seven (and limited to a maximum of nine) the rotorcraft must meet all the applicable paragraphs of this Part in effect on October 18, 1999, regardless of the original certification basis. Additionally, the amendment allows for an increase in maximum gross weight from 6,000 pounds up to 7,000 pounds for previously certified rotorcraft provided there is no increase in the maximum passenger-seat capacity beyond that for which the rotorcraft was certificated on October 18, 1999. If a passenger-seat capacity increase is requested in addition to an increase in gross weight above 6,000 pounds, then, once again, all applicable paragraphs of this Part in effect on October 18, 1999, must be met. b. Procedures. None. Page A - 4 ### Page 38 **CHAPTER 2. PART 27** **AIRWORTHINESS STANDARDS** **NORMAL CATEGORY ROTORCRAFT** **SUBPART B - FLIGHT** **GENERAL** AC 27.21. § 27.21 (Amendment 27-21) PROOF OF COMPLIANCE. a. Explanation. (1) This section provides a degree of latitude for the FAA/AUTHORITY test team in selecting the combination of tests or inspections required to demonstrate compliance with the regulations. Compliance should be shown for applicable combinations of gross weight, center of gravity, altitude, temperature, airspeed, rotor RPM, etc. Engineering tests are designed to investigate the overall capabilities and characteristics of the rotorcraft throughout its operational envelope. Testing will identify operating limitations, normal and emergency procedures, and performance information to be included in the FAA/AUTHORITY-approved portion of the flight manual. The testing must also provide a means of verifying that the rotorcraft’s actual performance, structural design parameters, propulsion components, and systems operations are consistent with all certification requirements. (2) Section 21.35 requires, in part, t hat the applicant show compliance with the applicable certification requirements, including flight test, prior to official FAA Type Inspection Authorization (TIA) testing. Compliance in most cases requires systematic flight testing by the applicant. After the applicant has submitted sufficient data to the FAA/AUTHORITY showing that compliance has been met, the FAA/AUTHORITY will conduct any inspections, flight, or ground tests required to verify the applicant’s test results. FAA/AUTHORITY compliance may be partially determined from tests conducted by the applicant if the configuration (conformity) of the rotorcraft can be verified. Compliance may be based on the applicant’s engineering data and a spot check or validation through FAA/AUTHORITY flight tests. The FAA/AUTHORITY testing should obtain validation at critical combinations of proposed flight variables if compliance cannot be inferred using engineering judgment from the combinations investigated. (3) Performance tests include mini mum operating speed (hover), takeoff and landing, climb, glide, height-velocity, and power available. Certain other performance tests, such as critical engine survey for multiengine installations, may be conducted to meet specific requirements. Detailed performance test procedures and allowable extrapolation or simulation limits are contained in the respective paragraphs in this AC. (i) Hover tests are conducted to determine various combinations of altitude, temperature, and gross weight for both in-ground-effect (IGE) and, if required Page B - 1 ### Page 39 by the applicant, out-of-ground effect (OGE) conditions. From these data, the hover ceiling may be calculated. (ii) Takeoff and landing tests are conducted to determine that a takeoff or landing can be safely executed without requiring exceptional piloting skill or favorable conditions at any approved combination of altitude, temperature, and gross weight. (iii) For rotorcraft other than he licopters, climb tests establish the variations of rate-of-climb at the best rate-of-climb or published climb airspeed(s) at various combinations of altitude, temperature, and gross weight. For helicopter, climb tests are conducted as required to determine the best rate-of-climb speed, Vy. (iv) Height-velocity tests are c onducted to determine the boundaries of the height versus airspeed envelope from which a safe landing can be accomplished following an engine failure. (v) Power available tests are conducted to verify the calculated installed specification engine performance model on which published performance is based. (4) The purpose of rotorcraft stability and control tests is to verify that the rotorcraft possesses the minimum qualitative and quantitative flying qualities and handling characteristics required by the applicable regulations. In order to assess the handling qualities, standardized test procedures must be utilized and the results analyzed by accepted methods. Section 27.21(a) allows calculation and inference which includes extrapolation and simulation, whereas § 27.21(b) requires demonstration of controllability, stability, and trim. Combinations of § 27.21(a) and (b) may be used to show compliance with the operating envelope limits. Test methods and equipment are described in individual paragraphs of this advisory circular. b. Procedures. (1) Efforts should begin early in the ce rtification program to provide advice and assistance to the applicant to ensure coverage of all certification requirements. The applicant should develop a comprehensive test plan which includes the required instrumentation. (2) The tests and findings specified in paragraph AC 27.21a(3) are required of the applicant to show basic airworthiness and probable compliance with the minimum requirements specified in the applicable regulations. After these basic findings have been submitted and reviewed, a Type Inspection Authorization, or equivalent, can be issued. The FAA/AUTHORITY will develop a systematic plan to spotcheck and confirm that compliance with the regulations has been shown. The test plan will consider combinations of weight, center of gravity, and RPM and cover the range of altitude and temperature for which certification is requested. Page B - 2 ### Page 40 AC 27.25. § 27.25 (Amendment 27-14) WEIGHT LIMITS. a. Explanation. (1) This section is definitive and spec ifies criteria for establishing maximum and minimum certificating weights. These weights may be based on those selected by the applicant, design requirements, or the limits for which compliance with all applicable flight requirements has been shown. (2) Typical requirements that may establish the maximum and minimum weight limits include: (i) Maximum: Structural limits, performance requirements, stability, and controllability requirements. (ii) Minimum: Autorotative ro tor RPM, stability, and controllability requirements. (3) Jettisonable External Cargo. (i) Section 27.25(c) was added by Amendment 27-11 to provide a basis for approving an increased gross weight that would be an external jettisonable load. Section 27.865, “External load attaching means, “ includes hoist and hook design features for the load attaching devices that were added to Part 27 but removed from § 133.43. Part 133, “Rotorcraft External-Load Operations,” was also amended (Amendment 133-5) concurrently to complement the changes to Parts 27 and 29. (ii) Approvals under the policy in Review Cases Nos. 37 and 55 of FAA Order 8110.6 were no longer necessary. These review cases concerned the policy/standards for external cargo configurations using a cargo hook whenever the standard limitations were exceeded. If the standard limitations were not exceeded, external cargo hooks and hoists and external cargo configuration approvals could be made under Part 133, Subpart D, prior to Amendment 133-5. (iii) In the preamble of Amendment 27-11 (Proposal 2-99, 41 FR 55454; December 20, 1976) the agency stated that “...§ 27.25(c) and § 29.25(c) are intended to provide only a total weight standard for approving the rotorcraft structure for operation under Part 133.” The policy in Review Case No. 55 also indicates the powerplant or propulsion system is also subject to evaluation for the increased weight. As indicated in § 27.865, fatigue substantiation of the external cargo attaching means is not required. The rotorcraft structure, rotors, etc., are only subject to fatigue evaluation under § 27.571 whenever the standard structural limitations are exceeded (Review Case No. 55). (iv) Whether or not the standard limitations are exceeded, the flight characteristics evaluations/standards of § 133.41 are appropriate even for engineering Page B - 3 ### Page 41 approval. Section 133.41 is also appropriate for the individual operator to obtain his operating certificate. The operator may use an FAA/AUTHORITY approved RFM supplement to prepare his own rotorcraft load combination flight manual required by § 133.47. b. Procedures. (1) It may not be possible to demons trate quantitatively all the flight requirements at the minimum weight because of test instrumentation requirements. The test team must ensure that the rotorcraft complies with the applicable requirements at the lowest permissible flying weight. This evaluation may be done qualitatively with the test instrumentation removed and with minimum crewmembers if no critical areas exist or are anticipated. Additionally, reasonable extrapolation is permitted. However, if critical areas at minimum flying weights are apparent, extrapolation should not be permitted. (2) Whenever a gross weight increas e under § 27.25(c) is requested, a TIA evaluation is necessary to evaluate the new limitations and ensure that § 133.41 for typical or representative cargo weights and/or shapes (or density) is satisfactory. All possible combinations of weights and shapes are not evaluated. The representative configurations may be noted in the RFM or RFM Supplement for the operator’s information. Sections 133.41 and 133.47 must be satisfied by the individual operator for the particular case at hand. The approved RFM or RFM Supplement should provide the necessary limitations and any other information about the representative cargo configurations evaluated. Section 133.41 also permits the operator to obtain approval of additional and unique cargo configurations provided approved limitations are observed. Paragraph AC 27.1581 concerns the RFM and its contents. (3) See AC 29-2C, Certificati on of Transport Category Rotorcraft, paragraph AC 29.571, concerning § 29.571, for fatigue substantiation and external cargo considerations that apply to § 27.571 as well. (4) Refer to AC 133-1A , Rotorcraft External-Load Operations in Accordance with FAR Part 133, for further information on airworthiness and flight manual policy for operators. AC 27.27. § 27.27 (Amendment 27-2) CENTER OF GRAVITY LIMITS. a. Explanation. (1) This regulation is definitive and requi res that the center of gravity limits be defined. Proof of compliance with all applicable flight requirements is required within the range of established CG’s. Along with the longitudinal CG limits, the lateral CG limits should either be established or determined to be not critical. Page B - 4 ### Page 42 (2) Ballast is usually carried during the flight test program to investigate the approved gross weight/center of gravity limits. Lead is the most commonly used form of ballast during rotorcraft flight testing although other types of ballast, such as water, may serve just as well. Water may have the added benefit of being jettisonable during critical flight test conditions. Care must be taken regarding the location of ballast. The strength of the supporting structures should be adequate to support such ballast during the flight loads that may be imposed during a particular test and for the ultimate inertia forces of § 29.561(b)(3). Of critical importance is the method of securing the ballast to the desired locations. To avoid any undesired in-flight movements of the ballast, a positive method of constraint is mandatory. The flight test crews should also visually verify the amount, location, and integrity of the ballast. The effects of mass moment of inertia on the flight characteristics due to the ballast locations should also be considered. The mass moment of inertia of the test rotorcraft should, to the extent possible, be the same as that expected in normal, approved loadings, especially during tests involving dynamic inputs. b. Procedures. (1) Center of gravity locations and lim its are of prime importance to rotorcraft stability and safety in flight. The primary concern is establishment of the longitudinal center of gravity limits. Lateral center of gravity limits with respect to longitudinal center of gravity limits are also important. The design of the rotorcraft is usually such that approximate lateral symmetry exists. This lateral symmetry can be upset by numerous probable lateral loadings possibly resulting in the necessity to establish lateral center of gravity limits. Stability and control characteristics may be seriously affected by loading outside the established center of gravity limits. The established center of gravity limits must be that as fuel is consumed, it is possible for the rotorcraft to remain within the established limits by acceptable loading and/or operating instructions. (2) Structural limits may restrict t he maximum forward longitudinal center of gravity limits. However, in most cases it is the maximum value established wherein adequate low speed control power exists to meet such requirements as § 27.143(c). Likewise, the maximum aft center of gravity limit may be a “structural limit,” but it usually is determined during flight test after the rotorcraft’s handling qualities tests have been conducted. Flight tests may reduce the “structural limit” CG envelope, but flight tests alone should not be used to expand the “structural limit.” Additional items which may influence the maximum aft center of gravity limits may be malfunctions of automatic stabilization equipment, excessive rotorcraft attitudes during critical phases of flight, or adequate control power to compensate for an engine failure. (3) Lateral center of gravity limits have become more critical because of the ever increasing utilization of the rotorcraft for such things as unusual and unsymmetric lateral loads, both internal and external. Maximum allowable lateral center of gravity limits have also influenced the results of the unusable fuel determination. Page B - 5 ### Page 43 (4) In summary, it is of prime import ance that longitudinal and lateral center of gravity limits be determined so that unsafe conditions do not exist within the approved altitude, airspeed, ambient temperature, gross weight, and rotor RPM ranges. All relevant malfunctions must be considered. AC 27.29. § 27.29 (Amendment 27-14) EMPTY WEIGHT AND CORRESPONDING CENTER OF GRAVITY. a. Explanation. The empty weight of the rotorcraft consists of the airframe, engines, and all items of operating equipment that have fixed locations and are permanently installed in the aircraft. It includes fixed ballast, unusable fuel, and full operating fluids except water intended for injection in the engines. (1) Fixed ballast refers to ballast t hat is made a permanent part of the rotorcraft as a means of controlling the empty weight CG. (2) Compliance with paragraph (b) of § 27.29 is accomplished by the use of an equipment list which defines the installed equipment at the time of weighing and the weight arm and moment of the equipment. b. Procedures. (1) Determination of the empty weight and corresponding center of gravity is primarily the responsibility of the manufacturing inspector. This determination is normally made on the production rotorcraft rather than the prototype. If the manufacturer wishes to avoid the necessity of weighing each production rotorcraft and he has been issued a production certificate, he may make a detailed proposal defining the procedures he will use to establish an empty weight and CG When his proposal is approved, he will weigh the first five to ten production rotorcraft and show that the rotorcraft will be within ±1 percent on empty weight and ±0.2 inches on CG After this procedure is established, the empty weight and CG may be computed except that at regular intervals, a rotorcraft will be weighed to ensure the tolerances are still being maintained; e.g., one in ten rotorcraft. (2) For prototype and modified rotorcra ft, it is only necessary to establish a known basic weight and CG position (by weighing) from which the extremes of weight and CG travel required by the test program may be calculated. See AC 91-23 (Pilots Weight and Balance Handbook) for a sample weight and balance procedure. (3) The weight and balance should be recalc ulated if a modification (or series of modifications) to the rotorcraft results in a significant change to the empty weight. Additionally, this change in empty weight should be reflected with the weight and balance information contained in the Rotorcraft Flight Manual (RFM) or Rotorcraft Flight Manual Supplement (RFMS). c. Ballast Loading and Type. Page B - 6 ### Page 44 (1) Ballast loading of the rotorcraft can be accomplished in any manner to achieve a specific CG location. It is acceptable for such ballast to be mounted outside the physical confines of the rotorcraft if the flight test objectives are not affected by this arrangement. In flight test work, loading problems will occasionally be encountered in which it will be difficult to obtain the desired CG limits. Such cases may require loading in engine compartments or other places not designed for load carrying. When this condition is necessary, care should be taken to ensure that local structural stresses are not exceeded or that the rotorcraft flight characteristics are not changed due to increased moments of inertia by attaching the ballast to extreme CG locations which may not be designed for the added weight. (2) There are basically two types of ba llast that may be used in loading. They are solids or liquid. The solids are usually high density materials such as lead while the liquid usually used is water. In critical tests, the ballast may be loaded in a manner so that disposal in flight can be accomplished. In any case, the load should be securely attached in its loaded position so shifting or interference with safety of flight will not result. AC 27.31. § 27.31 REMOVABLE BALLAST. a. Explanation. This regulation provides the option of using removable ballast to obtain desired center of gravity locations to determine compliance with the flight requirement of this Part. Fixed ballast used for flight operations after type certification must be documented in the type design data. Removable ballast is used primarily on small rotorcraft to control the CG with different passenger loadings although this regulation does not permit its use on transport rotorcraft. If removable ballast is used, the rotorcraft flight manual must include instructions regarding its use and limitations. b. Procedures. None AC 27.33. § 27.33 (Amendment 27-14) MAIN ROTOR SPEED AND PITCH LIMITS. a. Explanation. (1) General. This section requires the establishment of power-on and power-off main rotor speed limits and the requirements for low rotor speed warning. (2) Power-On. The power-on limits should be sufficient to maintain the rotor speed within these limits during any appropriate maneuver expected to be encountered in normal operations throughout the flight envelope for which certification is requested. In the past a minimum power-on range of approximately 3 percent has been required due to engine governor and engine operating characteristics. With the introduction of advanced engines and electronic engine controls, there may not be a need for a range. One fixed value may suffice. If substantiated, transient power-on values may also be acceptable. Page B - 7 ### Page 45 (3) Power-Off. The power-off rotor speed limits should be sufficient to encompass the rotor speeds encountered during normal autorotative maneuvers except for final landing phase (touchdown) for which rotor RPM may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. The limits should also be sufficient to cover the ranges of airspeed, weight, and altitudes for which certification is requested. It is not the intent of the rule to require the minimum and maximum limit values in conjunction with extremes such as maximum/minimum weights and/or high altitude. The minimum and maximum rotor speed requirements should be thoroughly evaluated at normal operating environment; i.e., at altitudes from approximately sea level to 10,000 feet, temperatures not at extremes, and weights as necessary for other tests and as required to readily establish the limit rotor speeds. Spot checks of the autorotative requirements should be made at the extremes of the flight envelope and environmental conditions during normal tests at those conditions. Under conditions where high autorotative rotor speeds may be encountered, it is acceptable for the pilot to adjust the controls to prevent overspeeding of the rotor. At light weight combined with low altitudes and extremely cold temperatures, the normal low pitch setting may not be sufficient to maintain autorotational rotor speed values within limits. If this occurs, the manufacturer may elect to adjust the low pitch stops as a maintenance procedure at extreme ambient conditions provided the flight and maintenance manuals clearly present the rigging requirements and procedures. There must be sufficient “overlap” of ambient conditions between configurations such that rerigging is not required whenever ambient temperature and surface elevation change slightly. Any downrigging of the low pitch stop must continue to ensure adequate clearance between controls and other rotorcraft structure and should be evaluated during flight test. Both the power-on and power-off limits may also be established by encountering critical flapping limits in some approved flight conditions such as high airspeed or sideward flight. (4) Low Speed Warning. If it is possible under expected operating conditions for the rotor speed to fall below the minimum approved values, the requirement exists for a low rotor speed warning. This warning is required on all single-engine rotorcraft and on multiengine rotorcraft where there is not an automatic increase in remaining engine(s) power output upon failure of an engine. Although not required by the rule, essentially all of today’s multiengine rotorcraft have a low rotor speed warning system installed. If the minimum power-on and power-off rotor speed limits are different, the warning signal should be at the higher speed, normally the power-on minimum rotor speed. One type of rotorcraft has a warning system cutout if the collective is full down, and other types have other warnings on the engine speed to indicate engine failure. All of these related warning systems must be evaluated with emphasis on ensuring adequate rotor speed. b. Determination and Testing. Refer to paragraph AC 27.1509 (§ 27.1509) for additional information on determining and testing rotor limits. Page B - 8 ### Page 46 **SUBPART B - FLIGHT** **PERFORMANCE** AC 27.45. § 27.45 (Amendment 27-21) PERFORMANCE--GENERAL. a. Explanation. (1) Section 45 of Part 27 lists some of the rules and standards under which the performance requirements are to be met. This paragraph will provide general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables which affect the flight test results that must be taken into account. Some of these variables are wind, temperature, altitude, humidity, rotorcraft weight, power, rotor RPM, center of gravity, etc. A thorough knowledge of the testing procedures and data reduction methods is essential and good engineering judgment must be used to determine acceptable test conditions. The test results should be analyzed and expanded by approved methodology within the guidelines of this paragraph. (2) Performance should be based on approved engine power as determined in paragraph b(5) below and not on any transient limits. Approved transient limits are basically for inadvertent overshoots of approved operational limits and any sustained operation in these transient limit areas usually requires some form of special maintenance. However, for such demonstrations as landing procedure demonstration and height-velocity (HV) determination, low rotor speeds (within approved limits) have been authorized. Such transients, if authorized, must be flight evaluated. (3) Where variations in the parameter on which a tolerance is allowed will have an appreciable effect on the test, the results should be corrected to the standard value of the parameter; otherwise, no correction is necessary. b. Procedures. (1) Winds for Testing. (i) Allowable wind conditions will vary with the type of test and will also be different for different types and gross weight rotorcraft. For example, higher winds can usually be tolerated for takeoff and landing tests than for hover performance. Higher winds can sometimes be tolerated during hover performance testing on rotorcraft with high rotor downwash velocities. Generally, unless the effects of wind on hover performance tests can be determined and/or accounted for, hover performance testing should be conducted in winds of 3 knots or less. Page B - 9 ### Page 47 (ii) In-ground-effect controll ability and maneuverability testing should be conducted in surface winds of less than 5 knots, or when higher steady wind conditions exist, with a maximum gust spread of 5 knots. (iii) As can be seen from the foregoi ng, there is no such thing as an exact allowable wind for a particular test or rotorcraft. The flight test team must decide on the allowable wind for each condition based on all available information and their engineering judgment. The following summary of allowable wind conditions is given for general guidance only: (A) Hover performance - 0 to 3 knots. (B) Height-velocity - 0 to 3 knots. (C) IGE controllability and maneuverability - 0 to 5 knots. (iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests. (2) Altitude Effects. Using FAA/AUTHORITY-approved methodology, hover performance may be extrapolated and/or interpolated from test data up to a maximum of ±4,000 feet. Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics should not be extrapolated higher than approximately 2,000 feet density altitude from the test altitude. Cruise stability/controllability tests should be evaluated at least at two different altitudes, the lowest practical altitude and approximately the highest cruise altitude requested for approval. This can allow an interpolation of approximately 10,000 feet. As in all testing, extrapolation and/or interpolation should only be considered if all available information and engineering judgment indicate that regulatory compliance can be met at the untested conditions. (3) Altitude Limitations. (i) Explanation. (A) Two altitudes are normally presented in the RFM to define the operating envelope of a rotorcraft; - Maximum operating altitude, and - Maximum takeoff and landing altitude. (B) Maximum operating altitude is an operating limitation required by § 27.1527 and delineates the maximum altitude to which operation is allowed. This altitude normally constitutes the maximum cruise or en route altitude. (C) Maximum takeoff and landing al titude is the hover in-ground-effect (IGE) ceiling for a rotorcraft as described in § 27.73. The hover ceiling and any Page B - 10 ### Page 48 information pertinent to takeoff and landing are presented in the performance information section of the RFM. For rotorcraft certified to CAR 6, Amendment 6-7 or any amendment of FAR 27, a hover ceiling may not be presented above the altitude at which H-V and IGE controllability tests were conducted plus allowable extrapolation, unless that extrapolated altitude is at least 7,000 feet. If the applicant elects to demonstrate these tests to an altitude below 7,000 feet, then that altitude is the maximum takeoff and landing altitude of the rotorcraft. The maximum takeoff and landing altitude may be coincident with, but never above the maximum operating altitude limitation. Takeoff and landing and hover ceiling data and presentation requirements are presented in §§ 27.51, 27.73 and 27.1587. (ii) Procedures. (A) In establishing the maximum takeoff and landing altitude, the following tests are normally required: ( 1) Takeoff (§27.51) ( 2) Climb (§§ 27.65 and 27.67) ( 3) Performance at minimum operating speed (§ 27.73) ( 4) Landing (§ 27.75) ( 5) Limiting height-speed envelope (§ 27.79) ( 6) IGE controllability (§ 27.143c) ( 7) Cooling (§§ 27.1041, 27.1043 and 27.1045) ( 8) Engine operating characteristics (§ 27.939) Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this AC. (B) As detailed in subparagraph b(2) above, the maximum allowable extrapolation of H-V, IGE controllability and engine operating characteristics is ±2,000 feet. Therefore, the maximum takeoff and landing altitude presented in the RFM is not normally more than 2,000 feet above the density altitude experienced at the high altitude test site, or for CAR 6, Amendment 6-7 and subsequent, unless test results were demonstrated to at least 7,000 feet. (C) If IGE controllability is demonstrated to at least 17 knots of wind at 7,000 feet, hover capability above 7,000 feet may be presented provided that the maximum demonstrated safe wind for takeoff and landing above 7,000 feet is specified in the RFM. Page B - 11 ### Page 49 (D) The requirements for data collection and presentation in the RFM vary depending upon the certification basis of the rotorcraft. These requirements are presented by regulation and amendment in figures AC 27.45-1 and AC 27.45-2. (E) The maximum takeoff and l anding altitude may be extrapolated no greater than the values given in paragraph b(2) and not above the lowest limiting altitude resulting from the requirements of subparagraph A of this paragraph. (4) Temperature Effects. (i) Background. (A) In the past, approved analyses were frequently accepted for determining the extreme temperature effects on performance and flight characteristics. With the introduction of newer, higher performance rotorcraft, advanced rotor blade designs, higher airspeeds, and higher blade tip Mach numbers, the previous methods have proven to be insufficient. Therefore, the performance and flight characteristics should be validated at extreme temperatures; however, analysis may be permitted if a suitable methodology is demonstrated. (B) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected by cold temperature in both the performance and flying qualities areas. Hot temperature conditions, although not shown to be as critical for flying qualities, should be given consideration. (C) Additionally, design deficiencies surfaced when the rotorcraft were exposed to temperature extremes and some of these difficulties were severe enough to require the redesign of equipment and/or materials. Therefore, to satisfy § 27.1309(a), the applicant needs to substantiate the total rotorcraft throughout the foreseeable range of operating temperatures. (ii) Procedures. (A) The FAA/AUTHORITY is responsible for verifying the effects of temperature on performance and handling characteristics. A limited flight verification, if necessary, could include spot checks of hover performance, IGE controllability, vibration, simulated power failure, static stability, height-velocity, VNE/VD evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations. (B) Extrapolation of test data should only be allowed if the applicant’s predicted or calculated data is verified by actual test, but in any case extreme caution should be used for extrapolations that are 10° C below or 20° C above those values tested. Page B - 12 ### Page 50 (5) Engine Power - Turboshaft Engine. (i) Background. (A) The purpose of rotorcraft performance flight testing is to obtain accurate quantitative flight test performance data to provide flight manual information. (B) Flight tests are designed to investigate the overall performance capabilities of the rotorcraft throughout its operating envelope. This testing furnishes information to be included in the flight manual and provides a means of validating the predicted performance of the rotorcraft with a minimum installed specification engine. (C) The power used to complete the flight manual performance must be based on power values no greater than that available from the minimum uninstalled specification engine after it is corrected for installation losses. A minimum uninstalled specification engine is one that, on a test stand under conditions specified by the engine manufacturer, will produce the certificated power at specification temperatures and/or speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specified power at a particular engine temperature or speed rating with higher allowable limits. The limit is the maximum value the installed engine is allowed in order to develop the specification power. Prior to installation of each engine in a rotorcraft, the performance is measured by the engine manufacturer. This is done by making a static test run in a test cell and referring the results to standard day, sea level conditions. The performance parameters obtained are presented as uninstalled engine characteristics on a test log sheet. This is commonly referred to as a “final run sheet.” Figure AC 27.45-3 compares a typical engine to one the manufacturer has certified as a minimum uninstalled certified engine. (D) After engine certification, t he engine manufacturer is responsible to ascertain that each engine delivered will produce, as a minimum, the certified power without exceeding specification operating values; therefore, a “final run sheet” is created for every engine produced. Additionally, if needed, arrangements can usually be made with the engine manufacturer to obtain a torque system calibration for individual engines. This will further optimize the accuracy of the engines used in the flight test program. The engine manufacturer will also provide predicted uninstalled power available for the various power ratings. This information may be derived from an engine computer “card deck” and from charts and tables in the engine detail installation manual. These data also provide engine performance for the range of altitudes and temperatures approved for the engine and include methods for correcting this performance for installation effects. The parameters contained in a typical “card deck” are plotted for one engine rating in figure AC 27.45-4. (E) Several power losses may be associated with installing an engine in a rotorcraft. Typical losses are air inlet losses, gear losses, air exhaust losses, and powered accessory losses such as electrical generators. Additional flight manual performance considerations are the torque indicating system accuracy and torque Page B - 13 ### Page 51 needle split. The predicted uninstalled power available engine characteristics cannot be assumed to be the actual power available after the engine is installed in the rotorcraft because this procedure would neglect the installation power losses. It is necessary to know the installation losses in order to determine the flight manual performance. Installation losses are reflected reductions in available power resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet and/or exhaust design. The rotorcraft manufacturer conducts tests to confirm the installed specification engine power available on which published performance is based. The specific methods used vary widely between manufacturers but usually include some combination of ground and flight tests. Figure AC 27.45-5 is a typical installed power available chart for one set of conditions. (F) The installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion may occur and there are changes in airflow routing. The rotorcraft manufacturer may elect to determine the installation losses for different flight conditions to take any airspeed advantages. This is acceptable if, for example, the hover performance is based on the actual power available from an installed minimum specification engine in a hover. Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual power available from an installed minimum specification engine at the published climb airspeed. This will allow the manufacturer to take advantage of, for example, increased inlet efficiency. (ii) Procedures. (A) The installed minimum specification engine power output has been predicted and calculated for various flight conditions. It is imperative that the predicted values be verified by actual flight test. The flight test involves obtaining engine performance measurements at various power settings, altitudes, and ambient temperatures. The data should be obtained at the actual flight condition for which the performance is to be presented (i.e., hover, climb, or cruise). (B) Following a power increas e, engine temperature and/or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM and/or temperature are held constant. This is a characteristic typical of turbine engines due largely to expansion of turbine blades and reduced clearances in the engine. Some engines may show a temperature increase at constant power due to engine or temperature sensing system peculiarities. An engine will usually establish a stabilized relationship of power parameters in approximately 2 or 3 minutes. For this reason, the following procedure should be used when obtaining in-flight engine data. ( 1) To determine the takeoff and 2 ½-minute values, first stabilize the engine at a low power setting. After stabilization, rapidly increase the power demand to takeoff and/or 2 ½-minute power levels. Record the engine parameters as soon as the specification torque, temperature, or speed is attained. Care must be taken not to Page B - 14 ### Page 52 exceed a limit. These readings should be obtained approximately 15 seconds after power is initially applied. ( 2) To determine the 30-minute and/or maximum continuous power values, approximately 2 to 3 minutes of stabilization time after power is increased is generally used, but up to 5 minutes stabilization time is allowed. The reason for the different procedures is when a pilot requires takeoff or 2 ½-minute power values he is in a critical flight condition and does not have the luxury of waiting for the engine(s) to produce rated power. Stabilization time is allowed for the maximum continuous and 30-minute ratings because these values are not associated with flight conditions for which power is needed immediately. (C) The in-flight measurements recorded with the engine(s) on the flight test rotorcraft must be corrected downward if the test engine is above minimum specification and corrected upward for a test engine that is below minimum specification. This correction is necessary to verify that a minimum specification engine installed on a production rotorcraft is capable of producing the power values used to compute the flight manual performance without exceeding any engine limit. In addition, if the production rotorcraft’s power measurement devices have significant (greater than 3 percent) power error, this error must be accounted for in a conservative manner. (D) On multiengine rotorcraft, the engine location may result in different installation losses between engines. If this condition exists, multiengine performance should be based on the total power available after considering the different installation losses and with minimum specification engines installed. One-engine-inoperative performance must be based on the loss of the engine which has the lowest installation losses. Additionally, the power losses due to such items as accessory bleed air, particle separators, engine driven accessories, etc., must be accounted for accordingly. (E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical NG speed to a constant corrected speed at cold temperatures. Others may limit power to a fuel flow value which would be encountered only at certain ambients. Others may limit by torque limiting devices. Therefore, power available data should be obtained at various ambients to verify that all limiting devices are functioning properly and have not been affected by the installation. (F) Through use, turbine engine power capabilities decrease with time. This is called engine deterioration. Deterioration is largely a function of the particular engine design, the manner, and the environment in which the engine is operated. There is a need, therefore, to provide a method which can be used in service to periodically determine the level of engine deterioration. A power assurance curve is usually provided to allow the flightcrew to know the power producing capabilities of any engine. A power assurance check is a check of the engine(s) which will determine that the engine(s) can produce the power required to achieve flight manual performance. This check does not have to be done at maximum engine power. Figure AC 27.45-6 is Page B - 15 ### Page 53 a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the rotorcraft flight manual performance. Some power assurance curves have maximum allowable NG limits that must not be exceeded for a given torque value. An in-flight power assurance check may be used in addition to the pretakeoff check. The validation of either check must be done by the methodology used to determine the installed minimum specification engine power available. For the in-flight power assurance check there must be full accountability for increased efficiency due to such items as inlet ram recovery, absence of exhaust reingestion, etc. A power assurance check done statically and one conducted in-flight must yield the same torque margin(s). An engine may pass power assurance at low power but still may not be capable of producing the rated power values. This occurs when the curve of corrected power and corrected temperature for the engine intersects the minimum uninstalled specification engine curve. If this condition exists, the entire power assurance and power available information may need to be reestablished. (6) Deteriorated Engine Power - Turboshaft Engine. (i) Background. (A) A specific engine model ma y have been certificated for operation with power which has “normally” deteriorated below specification. This “normal” deterioration refers to a gradual loss in engine performance, possibly caused by compressor erosion, as opposed to a sudden performance loss which may be due to mechanical damage. The application for deteriorated engine power should not be confused with the installed mechanical engine derating which is frequently used to match transmission and engine power capabilities. (B) The use of deteriorated power is intended to allow continued operations with an engine which is serviceable and structurally sound, although aircraft performance may be depreciated. The useful life of the engine may, therefore, be extended at a dollar savings to the operator. (C) Although installed performance is the primary topic in this discussion, considerations must be given to other operational characteristics and systems which may be affected by deteriorated engine power. These include: ( 1) Engine characteristics (§ 27.939). Surge margin, engine response, and air-restart capability might be affected and should be addressed, but flight testing may not be required depending on the individual engine/aircraft installation and fuel scheduling mechanism. ( 2) Performance of customer bleed air systems may be degraded slightly. No problem would be anticipated unless certain items within the system depend on a critical PC for their function. Page B - 16 ### Page 54 ( 3) The maximum attainable gas producer speed, and thus power available under certain ambients, may be affected if PC is an input to the fuel scheduling mechanism. ( 4) Systems for surge protection which schedule on P C such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The effect would normally be negligible unless when installed, the installation losses, combined with reduced PC because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings. (ii) Procedures. (A) The need for flight tests to verify predicted power available with deteriorated engines depends on the scope of testing which occurred during initial certification. If the original rotorcraft certification included flight testing as described in paragraph (5) (engine power-turboshaft engines) herein for validation of power available, the need for a demonstration with deteriorated engines is greatly diminished and perhaps eliminated. (B) If flight testing to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in paragraph (5) (engine power-turboshaft engines), except that the data would be corrected downward to a deteriorated engine runline. Efforts should concentrate on obtaining data in areas of the operational envelope where maximum gas producer speed is likely to be attained, or where bleed valves or other devices which schedule on gas producer discharge pressure are likely to function. On many installations maximum gas producer speed will occur with cold temperatures and high altitudes; bleed valves and other devices which schedule on gas producer discharge pressure are most likely to function and reduce power on a hot day at low altitude. (C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated engine monitoring procedures will be introduced when allowing the use of deteriorated power; for example, an in-flight trend monitoring program with the associated bookkeeping duties may be required. Such an in-flight procedure must be evaluated by flight tests as described in paragraph (5) (engine power-turboshaft engines) herein. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure already in place for the specification engine, which could eliminate the need for flight test evaluation. Page B - 17 ### Page 55 AC 27.45A. § 27.45 (Amendment 27-21) GENERAL. a. Explanation. Amendment 27-21 adds § 27.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered rotorcraft. This information is to provide the pilot a means of determining, prior to takeoff, that each engine will produce the power necessary to achieve the performance presented in the rotorcraft flight manual (RFM). b. Procedures. All of the policy material pertaining to this section remains in effect. In addition, the power assurance information included in the RFM should be verified. Although this requirement is normally met with a power assurance curve, other methods of compliance may be proposed. AC 27.45B. § 27.45 (Amendment 27-29) GENERAL. a. Explanation. Amendment 27-29 added the requirements for certification of 30-second/2-minute One Engine Inoperative (OEI) power ratings. For rotorcraft approved for the use of 30-second/2-minute OEI, partial power checks currently accomplished with approved power assurance procedures for lower power levels may not be sufficient to guarantee the ability to achieve the 30-second power level. b. Procedures. Information provided in Appendix 1 of this AC includes guidance material on power assurance procedures to assure that the OEI power level can be achieved Page B - 18 ### Page 56 **CERTIFICATION BASIS** . FAR 27 CAR 6 CAR 06 Rqmts 27-Amdt. 21 27-Amdt. 2 Original 6-Amdt. 7 6-Amdt. 4 Original Original H-V Ref. 27.25 27.79 27.1519 27.1587 6.116 6.741 6.743 Test Conditions 1. MGW Sea Level 2. Max. OGE wt. Lesser of: a. Max. alt. cap. b. 7000' Hd 1. MGW Sea Level 2. Wt. selected by applicant at lessor of: a. Max. alt. cap. b. 7000' Hd 1. MGW Sea Level 2. Wt. selected by applicant at lessor of: a. Max. alt. cap. b. 7000' Hd 1. MGW Sea Level 2. Wt. selected by applicant at lessor of: a. Max. alt. cap. b. 7000' Hd 1. MGW Sea Level 2. Wt. selected by applicant max. t.o. and ldg. alt. 1. MGW Sea Level 2. Wt. selected by applicant at max. t.o. and ldg. alt. 1. No specific wt. and alt. requirements. RFM 3. H-V is perf. Info. 4. Max. alt. For which H-V is valid 3. H-V is perf. info. 4. If H-V wt is less than IGE wt., wt. becomes limitation for t.o./ldg. 3. H-V is perf. info. 3. H-V is perf. info. 3. H-V is perf. info. 3. H-V is operating limitation. H-V is operating limitation. AC 27- 1B Para AC 27.45 & AC 27.51 Remarks 5. If H-V is less than OGE wt., H-V wt. becomes limit. 6. Applicant is encouraged to demo H-V to WAT limits. 7. Hover data may be shown above 7000' if H-V & IGE are demo'd to 7000' FIGURE AC 27.45-1 H-V Requirements Page B - 19 ### Page 57 **CERTIFICATION BASIS** FAR 27 CAR 06 Rqmts 27-Amdt. 21 Original Original Original IGE CONTROL Ref. 27.25 27.143 27.1587 AC 27-1B Para AC 27.45 & AC 27.143 Test Conditions 1. MGW Sea Level 2. Max. IGE wt. Lesser of: a. Max. alt. cap. b. 7000' Hd 3. Critical CG 4. Critical Rotor RPM 5. Wind of not less than 17 kts. 1. MGW Sea Level 2. Wt. selected by applicant to max. t.o. and ldg. alt. 3. Critical CG 4. Critical Rotor RPM 5. Critical wt. 6. Wind of not less than 20 mph 1. MGW Sea Level 2. Max. approved wt. for t.o./ldg at alt above sea level. 3. Critical CG 4. Critical Rotor RPM 5. Wind not less than 20 mph. 1. No specific requirement. RFM 6. Max. safe wind is perf. info. 7. Max. safe wind is perf. info. 6. Max. safe wind is perf. info. Remarks 7. If 17 kts. wind demo'd to alt. less than 7000', a corresponding WAT limit must be established. **FIGURE AC 27.45-2 IGE CONTROLLABILITY REQUIREMENTS** Page B - 20 ### Page 58 (OCR) **600 500 CORRECTED SHAFT HORSEPOWER ENG S/N 12345 400 300 MIN CERTIFIED** SPECIFICATION ENGINE 200 100 1000 1100 1200 1300 1400 1500 CORRECTED TOT °F FIGURE AC 27.45-3 SHAFT HORSEPOWER VS TURBINE OUTLET TEMPERATURE - SEA LEVEL STANDARD DAY Page B - 21 ### Page 59 (OCR) 20 -10° -30° to -50° 50° 16 PRESSURE ALTITUDE - FEET (X 1000) O L 200 300 400 500 SHAFT HORSEPOWER AVAILABLE FIGURE AC 27.45-4 UNINSTALLED TAKEOFF POWER AVAILABLE Page B - 22 ### Page 60 (OCR) 20 18 ·0 > 10-501 C -20 ° 0 14 TRANSMISSION LIMIT PRESSURE ALTITUDE - FEET (X 1000) 200 300 400 SHAFT HORSEPOWER AVAILABLE FIGURE AC 27.45-5 INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM Page B - 23 ### Page 61 (OCR) CECCE EXAMPLE: OAT = 17.5°C TOT = 760° C Hp = 4000 FT REQUIRED TORQUE = 70% 80 100 PERCENT TORQUE FIGURE AC 27 45-6 POWER ASSURANCE CHECK CHART ZERO AIRSPEED GENERATOR LOAD 100 AMPS 40 CONDITIONS: RPM 100% ANTI-ICE OFF 10 800°C 780°C 760°C 740°C 720°C 700°C 680°C 660°C Page B - 24 ### Page 62 AC 27.51. § 27.51 TAKEOFF. a. Explanation . Section 27.51 details the conditions under which takeoff data must be obtained. The flight manual must contain the technique(s) to be used to obtain the published flight manual takeoff procedures. Technique should not be confused with exceptional pilot skill and/or alertness as mentioned in § 27.51. Because rotorcraft differ, different pilot techniques are sometimes required to achieve the safest and most optimum takeoff performance. The recommended technique that is published in the flight manual must be determined to be one that the operational pilot can duplicate using the minimum amount of type design cockpit instrumentation and the minimum crew. Only rotorcraft takeoff techniques will be covered in this section. b. Background . (1) Certain special takeoff techniques are necessary when a rotorcraft is unable to take off vertically because of altitude, weight, power effects, or operational limitations. The recommended technique used to take off under such conditions is to accelerate the rotorcraft in-ground-effect (IGE) to a predetermined airspeed prior to climbout. Takeoff tests are performed to determine the best repeatable technique(s) for a particular rotorcraft over the range of weight and altitude for which certification is requested. (2) Utilizing the total power available to execute a takeoff may not be operationally feasible due to such items as HV or aircraft attitude constraints. In such situations, hover power required plus some power increment may be the maximum recommended for use. (3) Wheel or skid height shoul d be not less than that demonstrated satisfactorily for the high speed, low altitude portion of the HV curve, or that height below which ground contact may occur when accomplishing takeoff procedures. (4) For rotorcraft fitted with wheel s, a running takeoff procedure may be accepted. c. Procedures . (1) There are different takeoff profiles which may be used to complete a maximum performance takeoff in a rotorcraft. The manufacturer will normally determine which method is best for a particular rotorcraft. The most commonly accepted method is the hover and level acceleration technique. In this technique, the rotorcraft is stabilized in a hover at the reference height. From the stabilized hover, the rotorcraft is accelerated to the climbout airspeed using the predetermined takeoff power. When the desired climbout airspeed is achieved, the rotorcraft is rotated and the climbout is accomplished at the scheduled airspeed(s) and constant rotor RPM. Power adjustments may be accomplished to maintain the targeted power except where procedure requires high workload outside the cockpit (i.e., that portion of takeoff where horizontal acceleration close to the ground has pilot scan outside the cockpit and Page B - 25 ### Page 63 adjustment of engine torque or temperature would require an undue increase in workload). The recommended takeoff procedure must be demonstrated to remain clear of the HV “avoid” areas without requiring exceptional piloting skill or exceptionally favorable conditions. (2) The hover reference height is es tablished as the minimum skid or wheel height above the takeoff surface from which a takeoff can consistently be accomplished in zero wind without contacting the runway surface. The takeoff must be accomplished with power fixed at the power required to hover at the hover reference height and must not require exceptional piloting skill to avoid runway surface contact. AC 27.65. § 27.65 (Amendment 27-14) CLIMB: ALL ENGINES OPERATING. a. Explanation . (1) Rotorcraft other than helicopters. (i) Section 27.65 requires that the steady rate of climb be determined for each rotorcraft other than helicopter with maximum continuous power on each engine for the range of weights, altitudes, and temperatures for which certification is requested. Equivalent levels of safety have been found wherein the applicant was allowed to select a climb airspeed that was not the actual VY. The selected airspeed must be consistent with the speed used to show compliance with such items as cooling, stability, etc. The rate of climb resulting from the selected climb airspeed versus that from the actual VY shall not differ to an extent that a pilot will be encouraged, by appreciable increases in climb performance, to fly a climb airspeed different from that published in the flight manual. (ii) For rotorcraft other than he licopters, the climb performance data obtained above must be used to show that a minimum climb gradient can be achieved for each weight, altitude, and temperature within the range for which certification is required. This gradient must be at least 1:10 if testing is done to determine the required takeoff distance over a 50-foot obstacle. If this option is selected, an explanation of the takeoff distance determination requirements and procedures may be found in paragraph AC 29.63 of AC 29-2C. (iii) If takeoff distance is not det ermined, the minimum climb gradient must be 1:6 for standard sea level conditions. (2) For helicopters, V Y must be determined for standard sea level conditions at maximum weight using maximum continuous power on each engine. Although not required, the steady rate of climb may be determined using the procedure in paragraph AC 27.65c of this section (Procedure to Determine All-Engine-Operating Climb Performance). Page B - 26 ### Page 64 (3) For helicopters, if V NE at any altitude is less than the maximum gross weight sea level standard day condition VY, the steady rate of climb must be determined at the climb speed(s) selected by the applicant not to exceed VNE. The climb performance must be determined from 2,000 feet below the altitude from where VNE intersects VY up to the maximum altitude for which certification is requested. This should be done utilizing maximum continuous power on each engine with the landing gear retracted. b. Procedure to Determine V Y. (1) Sawtooth climbs may be used to determine V Y. If such a technique is used, climbs should be flown in pairs on opposite headings 90° to the wind at the test altitude. This procedure will minimize any windshear effects. All testing must be done in smooth air. Windshear is usually an indication of unstable air or a temperature inversion and must also be avoided. The climbs are flown on reciprocal headings for approximately 5 minutes or through an altitude band using maximum continuous power at a constant airspeed. Periodic power adjustments may be necessary. Additional reciprocal heading climbs must also be conducted at different airspeeds above and below the airspeed at the lowest point of the power required versus airspeed curve. This technique can be repeated at different altitudes to obtain VY throughout the altitude range. (2) Level flight performance (speed pow er) may also be used to determine VY. The testing should be done in smooth air. The advantage of this method is that less time is required, and the accuracy is equivalent to the sawtooth climb method. The test can be repeated at various altitudes to determine the VY throughout the altitude range desired for the rotorcraft. The test at each altitude should be conducted at a constant weight over sigma (W/σ). The test is normally started at the desired W/σ with maximum continuous power, or at VNE, in level flight. A series of points should be taken, reducing airspeed 10 to 15 knots between points, with the lowest speed point around 20 to 30 knots. Weight should be computed for each point and the test altitude adjusted to maintain a constant W/σ. After the data are reduced to standard day conditions, the minimum power required airspeed will be the VY speed. (3) Prior to the flight test, the roto rcraft should be ballasted to the desired gross weight and the critical center of gravity. The airspeed should be stabilized prior to data acquisition. Data to be recorded includes time, altitude, airspeed, ambient temperature, engine parameters, torque(s), rotor RPM, fuel reading, aircraft heading, external configuration, etc. Power setting, weight, and climb airspeed should be planned prior to flight. For some turboshaft engines, temperature and/or engine speed limits may be reached prior to a limiting torque. The test team should verify that the resulting power utilized in these tests closely approximates the power producing capabilities of a minimum installed specification engine. c. Procedure to Determine All-E ngine-Operating Climb Performance. Page B - 27 ### Page 65 (1) Background . Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to validate the rotorcraft’s climb performance. By-products are a qualitative evaluation of the rotorcraft handling characteristics in a climb and engine data to assist in the determination of installed power available. (2) Techniques . The climbs are conducted on reciprocal headings at the established airspeed(s) through the target altitude range. The same parameters are recorded as during sawtooth climbs. The rotorcraft will usually climb very rapidly during the first few thousand feet; therefore, the data acquisition method must be timely if accurate results are expected. This procedure is usually repeated at weight extremes. The resulting data must then be corrected for power and weight. Power and weight corrections are satisfactory, provided the test powers and weights closely approximate the target values to make the weight and power corrections small. Once this data is finalized and corrected for all the flight test variables, interpolation for intermediate weights can be made with a high degree of reliability. If the rotorcraft has any stability augmentation system, vent systems, etc., which may influence the climb performance, then it must be accounted for. Caution should be taken that anti-ice, air-conditioning, etc., are not on unless the performance is being established specifically for those conditions. AC 27.65A. § 27.65 (Amendment 27-33) CLIMB: ALL ENGINES OPERATING. a. Explanation . Amendment 27-33 added the requirement to determine the steady rate of climb, for helicopters, from sea level up to the maximum altitude for which certification is requested. Although not specifically stated in the rule, the rate of climb should be determined at VY or, if VNE at any altitude is less than the maximum gross weight sea level standard day condition VY, the steady rate of climb at these altitudes must be determined at a climb speed(s) selected by the applicant not to exceed VNE. b. Procedures . The policy material pertaining to the procedures outlined in this section remain in effect. AC 27.67. § 27.67 (Amendment 27-23) CLIMB: ONE ENGINE INOPERATIVE. a. Explanation . (1) Section 27.67 requires that for multiengine normal category rotorcraft, the steady rate of climb or descent with one engine inoperative must be determined at VY (or at the speed for minimum rate of descent) for maximum gross weight. (2) The rate of climb (or descent) will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or 30-minute minimum specification installed power available values. The landing gear should be retracted if it is retractable. b. Procedures . Page B - 28 ### Page 66 (1) The procedure discussed in paragr aph AC 27.65 for all-engines-operating climb performance is also applicable to the OEI condition. For twin-engine rotorcraft that are shown not to have a “critical engine” with respect to performance characteristics, both engines may be used to simulate the appropriate single-engine power available during these tests. (2) Adequate testing must be accomp lished to determine the rotorcraft’s OEI climb performance at maximum gross weight for all variations in altitude and temperature for inclusion in the Rotorcraft Flight Manual. AC 27.71. § 27.71 (Amendment 27-21) GLIDE PERFORMANCE. a. Explanation . (1) Performance capabilities during st abilized autorotative descent are useful tools to assist the pilot when all engines fail. This information is also useful in determining the suitability of available landing areas along a given route segment. (2) Two speeds are of particular impor tance, the speed for minimum rate of descent and the speed for best angle of glide. These speeds along with glide distance information are required as flight manual entries per § 27.1587. The speed for minimum rate of descent is useful for engine failure conditions at higher altitudes and the pilot is required to perform some time-related task, engine restart, float inflation, radio calls, etc. The speed for best angle of glide is a somewhat higher speed that is of particular use when it is necessary to reach a distant landing area. These speeds, when utilized in conjunction with appropriate rotor RPM and glide angle (or rate of descent) can be used to calculate the maximum horizontal distance available from a particular altitude assuming zero wind conditions. (3) A third speed, recommended autor otation speed, may be provided in addition to minimum rate of descent speed and maximum glide angle speed. The recommended speed for autorotation is usually optimized to assure an effective flare capability and yet be slow enough to allow a controlled, relatively slow touchdown condition. Recommended autorotation speed is ordinarily between the minimum rate of descent and maximum glide angle speeds. The recommended autorotation speed may be provided in the Rotorcraft Flight Manual. The relationship between minimum rate of descent, best glide angle, and recommended autorotation speed is shown in figure AC 27.71-1. (4) Forward center of gravity is usua lly critical; however, center of gravity effects should be spot-checked to confirm this for a given design. b. Procedures . Page B - 29 ### Page 67 (1) Tests are conducted at speeds which bracket the anticipated speeds for minimum rate of descent and best glide angle. On a power required plot, the speed for minimum power required approximates the speed for minimum rate of descent. The speed for maximum range glide may be estimated by drawing a tangent from the origin to the power required curve. (2) Autorotative performance tests may be conducted in conjunction with the climb performance tests. The required data are similar for both tests and it is sometimes convenient and efficient to run alternating climbs and descents through a desired altitude band. Descents should be conducted on reciprocal headings and results averaged in the same manner as climb performance tests. (3) A reduction in rotor RPM from the normal power-on value may enhance autorotative performance. If the applicant wishes to develop autorotative performance at RPM values significantly below the governing or power-on range, the practicality of reducing and controlling RPM at the lower value and of then increasing RPM as a landing is approached, must be considered. At low weights and low density altitudes, full down collective may automatically produce lower RPM values and this condition is, of course, acceptable provided the approved power-off RPM range is not exceeded. (4) During autorotation tests, care must be taken to make certain that no engine power is delivered to the rotor drive system since a very small amount of power can have a large effect on descent performance. Page B - 30 ### Page 68 (OCR) **RATE OF** **DESCENT RECOMMENDED** **AUTOROTATIONAL SPEED** **BEST GLIDE ANGLE** **MINIMUM** **RATE OF DESCENT T** AIRSPEED FIGURE AC 27.71-1 AUTOROTATIONAL CHARACTERISTICS - TYPICAL Page B - 31 ### Page 69 AC 27.73. § 27.73 PERFORMANCE AT MINIMUM OPERATING SPEED. a. Explanation . (1) The word “hover” applies to a rotorc raft that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air. For the purpose of this manual, the word “hover” will mean pure hover. (2) The regulatory requirement for hover performance, § 27.73, refers to hover in ground effect (IGE). For some applications, such as external load operations, hover performance out-of-ground effect (OGE) is necessary; however, it is not required by this section. Hover OGE is that condition, where an increase in height above the ground will not require additional power to hover. Hover OGE is the absence of measurable ground effect. It can be less than one rotor diameter at low gross weight increasing significantly at high gross weight. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle one and one-half rotor diameters above the surface. (3) The objective of hover performance tests is to determine the power required to hover at different gross weights, ambient temperatures, and pressure altitudes. Using nondimensional power coefficients (Cp) and thrust coefficients (Ct) for normalizing and presenting test results minimizes the amount of data required to cover the rotorcraft’s operating envelope. (4) Hover performance tests must be conducted over a sufficient range of pressure altitudes and weights to cover the approved ranges of those variables for takeoff and landing. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects. (5) The hover ceiling for which data should be obtained and subsequently presented in the flight manual should be the same height consistent with the minimum hover height demonstrated during the takeoff tests. Refer to paragraph AC 27.51 for the procedure to determine this hover height. b. Procedures. (1) Two methods of acquiring hover performance data are the tethered and the free flight techniques. The tethered technique is accomplished by tethering the rotorcraft to the ground using a cable and load cell. The load cell and cable are attached to the ground tie-down and to the rotorcraft cargo hook. The load cell is used to measure the rotorcraft’s pull on the cable. Hover heights are based on skid or wheel height above the ground. During tethered hover tests, the rotorcraft should be at light gross weight. The rotorcraft will be stabilized at a fixed power setting and rotor speed at the appropriate skid or wheel height. Once the required data are obtained, power should be varied from the minimum to the maximum allowed at various rotor RPM. This technique will produce a large Ct/Cp spread. The load cell reading is recorded for each Page B - 32 ### Page 70 stabilized point. The total thrust the rotor produces is equal to the rotorcraft’s gross weight plus the weight of the cables and load cell plus cable tension. Care must be taken that the cable tension does not exceed the cargo hook limit or load capacity of the tie-down. For some rotorcraft, it may be necessary to ballast the rotorcraft to a heavy weight in order to record high power hover data. (2) The pilot maintains the rotorcraft in position so that the cables and load cell are perpendicular to the ground. To ensure the cable is vertical, two outside observers, one forward of the rotorcraft and one to one side, can be used. Either hand signals or radio can be used to direct the pilot. The observers should be provided with protective equipment. Positioning can also be accomplished by attaching two accelerometers to the load cell which sense angle or movement along the longitudinal and lateral axes. Any displacement of the load cell will be reflected on instrumentation in the cockpit, and by reference to this instrumentation, the rotorcraft can be maintained in the correct position. Increased caution should be utilized as tethered hover heights are decreased because the rotorcraft may become more difficult to control precisely. The tethered hover technique is especially useful for OGE hover performance data because the rotorcraft’s internal weight is low and the cable and load cell can be jettisoned in the event of an engine failure or other emergency. (3) To obtain consistent data, the wi nd velocity should be less than 3 knots as there are no accurate methods of correcting hover data for wind effects. Rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are: (i) Engine torque. (ii) Rotor speed. (iii) Ambient temperatures. (iv) Pressure altitude. (v) Fuel used (or remaining). (vi) Load cell reading. (vii) Generator(s) load. (viii) Wind speed and direction. As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in paragraph AC 27.73b(4) will determine if any problems, Page B - 33 ### Page 71 such as load cell malfunctions, have occurred. The free flight hover data must fall within the allowable scatter of the tethered data. (4) If there are no provisions or equipm ent to conduct tethered hover tests, the free flight technique is also a valid method. The disadvantage of this technique as the primary source of data acquisition is that it is very time consuming. In addition a certain element of safety is lost OGE in the event of an emergency. The rotorcraft must be reballasted to different weights to allow the maximum Ct/Cp spread. When using the free flight technique, either as a primary data source or to substantiate the tethered technique, the same considerations for wind, recorded parameters, etc., as used in the tethered technique apply. Free flight hover tests should be conducted at CG extremes to verify any CG effects. If the rotorcraft has any stability augmentation system which may influence hover performance, it must be accounted for. (5) It is extremely diffi cult to determine when a rotorcraft is hovering OGE at high altitudes above ground level since there is no ground reference. In a true hover, the rotorcraft will drift with the wind. Numerous techniques have been tried to allow OGE hover data acquisition at high altitudes, all of which have resulted in much data scatter. Until a method is proposed and found acceptable to the FAA/AUTHORITY, OGE hover data must be obtained at the various altitude sites where IGE hover data are obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet. AC 27.75. § 27.75 (Am endment 27-14) LANDING. a. Explanation . (1) This rule incorporates all of t he landing requirements for Part 27 rotorcraft. (2) As with other flight maneuver s, landings must be accomplished with acceptable flight and ground characteristics using normal pilot skills. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 27.45. As in other performance areas, engines must be operated within approved limits. (3) Landing. Approach and landing pat h requirements are stated in general terms in paragraphs (a)(1) and (a)(2) of § 27.75. The approach path must allow smooth transition for a one-engine-inoperative landing and adequate clearance from potentially hazardous HV combinations. (4) All-engine-out landing. Section 27.75(b) contains the certification requirement for “last” engine failure and all-engines-inoperative landing. The rule states that it must be possible to make a safe landing after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously, although complete power failure has occurred in twin-engine rotorcraft with Category A engine isolation. This requirement assures that in the event of cockpit mismanagement, fuel Page B - 34 ### Page 72 exhaustion, improper maintenance, fuel contamination, or unforeseen mechanical failures, a safe autorotation entry can be made and a safe power-off landing can be affected. Two separate aspects of this rule are normally evaluated at different times during the test program. The “last” engine failure is normally evaluated during cruise or VNE engine failure testing where instrumentation and critical loading have been established for those test conditions. The all-engine-out landing is ordinarily conducted in conjunction with an HV or landing distance phase where ground instrumentation and safety equipment are available. b. Procedures . (1) Instrumentation/Equipment. Ai rcraft instrumentation may include engine and flight parameters, control positions, power lever position, and landing gear loads. A record of rotor RPM at touchdown is necessary to assure it does not exceed transient limits. Rotor RPM at touchdown may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. A crash recovery team with the support of a fire engine is highly desirable. (2) The one-engine-inoperative landing is si milar in many respects to the HV tests described in paragraph AC 27.79. Most of the comments, cautions, and techniques for HV also apply here even though the typical flight conditions are less critical than limiting HV points due to a lower power level and an established rate of descent. The approach is made at a predetermined speed with one engine inoperative. The speed is reduced and the rotorcraft is flared to a conventional one-engine-inoperative landing. (3) Power. Power should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of engine topping to minimum specification values for the range of atmospheric variables to be approved. This is frequently done by installing an adjustable device in the throttle linkage with a control in the cockpit so that engine topping can be accurately adjusted for varying ambient conditions. With such a device in the control system it becomes vitally important to check topping power prior to each test sequence. (4) Aircraft Loading. Aft center of gr avity is usually most critical because visibility constraints limit the degree to which the pilot can see the landing surface during the flare. If a weight effect is shown, a minimum of two weights should be flown at each test altitude. One weight should be the maximum weight for prevailing conditions, and the other should provide a sufficient spread to validate weight accountability. (5) All-engine-out landing. (i) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion of this paragraph, § 27.75(b) contains two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. The second Page B - 35 ### Page 73 aspect of this rule requires that a landing from autorotation be possible. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. If a complete company test program has documented an all-engine-out landing to the GW/ (gross weight/density ratio) limit, verification tests may be initiated at those limiting weight conditions. If not, buildup testing should be initiated at light weight. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings-one at sea level and one near maximum takeoff and landing altitude. (ii) Demonstrated compliance with th is requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to a reasonable value (less than 40 KTAS is recommended) to ensure a reasonable chance of survivability for the all engine failure condition. On multiengine rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to this relatively low inertia level, considerable collective may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which may lead to rotorcraft damage in testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines cut off. (iii) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train. Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished. Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a multiengine rotorcraft totally without power. The all-engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests. (6) Prior to conducting these tests, t he crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons idle power be used if a “needle split” (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure Page B - 36 ### Page 74 needle split is attained. In other cases a temporary detent between idle and cutoff was used on the throttle. In a third case the engine was actually shut down on sample runs to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as is reasonable to dissipate speed and build RPM Rotor RPM must stay within allowable limits. Aft center of gravity is ordinarily critical due to visibility and flarability. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it must be within allowable structural limits. AC 27.79. § 27.79 (Amendment 27-14) LIMITING HEIGHT-SPEED ENVELOPE. a. Explanation . (1) The height-speed envelope is normally referred to as the height-velocity (HV) diagram. It defines an envelope of airspeed and height above the ground from which a safe power-off or OEI landing cannot be made. The diagram normally consists of three portions: (a) the level flight (cruise) portion, (b) the takeoff portion, and (c) the high speed portion. See figure AC 27.79-1. The high speed portion is omitted on occasions when it can be shown that the rotorcraft can suffer an engine failure at low altitude and high speed (up to VH) and make a successful landing or climb out on the remaining engine(s). (2) Power failure, engine fa ilure, throttle chop, or other similar terms used in this discussion mean a simulated engine failure. The actual shutdown of an engine to simulate an engine failure should not be necessary if the simulated procedure ensures that the engine power is suddenly removed from driving the rotor and remains so. The normal fuel control deceleration schedule is usually satisfactory for the power removal for turbine engines but the flight/ground idle speed may have to be set lower than normal for HV testing. (3) The avoid areas of the HV diagram are separated by the takeoff corridor. This corridor should be wide enough to consistently permit a takeoff flight path clear of the HV diagram using normal pilot skill. The takeoff corridor should always permit a minimum of ±5 knots clearance from critical portions of the diagram. (4) The knee of the curve separates t he takeoff portion from the cruise portion and is defined as the highest speed point on the low speed portion of the HV envelope. Altitudes above this point are considered cruise, or “fly-in,” points, and these test points require a minimum time delay of 1 second between throttle chop and control actuation (reference § 27.143(d)). Altitudes below the knee represent takeoff profile points. For test points in the takeoff portion, takeoff power (or a lower power selected by the applicant as an operating procedure) and normal pilot reaction time for corrective control actuation will be used. (5) Since the HV diagram may repr esent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer’s Page B - 37 ### Page 75 buildup program should be reviewed to determine the amount of conservatism in the HV diagram (if any). It should be remembered that the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points are approached from above and below respectively. Portions near the knee are initially evaluated at high speed with subsequent backing down of the speed. In most rotorcraft the effective flare airspeed is critical. At airspeeds slightly below this value, the ability to arrest and control descent rates through use of an aft cyclic flare may be greatly diminished. Extreme care should be exercised when “backing down” to lower speeds. (6) In addition to the on-board and gr ound instrumentation, a motion picture camera or other position measuring equipment should cover each run. (7) For FAA/AUTHORITY tests, the minimum required crew and the minimum instrument panel display presented for certification should be used. Ground safety equipment should be provided. (8) This test is the least predictable of all the performance items. Therefore, the expansion and extrapolation of test data are questionable. Weight may not be extrapolated to higher values. In order to extrapolate HV data to higher altitudes, any analytical method must have FAA/AUTHORITY approval. In lieu of pure analytical methods, simulations have been used successfully, especially for multiengine rotorcraft. In either case, the maximum allowable extrapolation should be limited to 2,000 feet density altitude (HD). HV test weights for normal category rotorcraft are the maximum weight at sea level and some lessor weight at high density altitudes. The high density altitude HV curve needs to be defined only to 7,000 feet and may be a lower altitude if the rotorcraft does not have the performance capabilities to attain 7,000 feet. A weight less than the maximum weight may be used to define the high density altitude HV curve, but this weight should not be less than the maximum weight that will allow hovering out-of-ground effect. For a given diagram, typical weight reductions that are necessary as altitude is increased can be conservatively estimated by maintaining a constant gross weight divided by density ratio, GW/σ. See figure AC 27.79-2, Part A. If weight is not varied, an enlarged HV diagram is required for safe power-off landing as density altitude is increased. See figure AC 27.79-2, Part B. Another method of presentation is to show varying weights at a constant density altitude. (See figure AC 27.79-2, Part C.) (9) Vertical takeoff and landing (VTO L) testing normally does not require separate HV testing. The takeoff and landing tests take on the combined characteristics of takeoff, landing, and HV tests. b. Procedures . (1) Instrumentation. Page B - 38 ### Page 76 (i) Ground Station . The ground station must have equipment and instrumentation to determine wind direction and velocity, outside air temperature, and if the test rotorcraft has reciprocating engines, humidity. Since the tests must be conducted in winds of 2 knots or less, a smoke generator is highly recommended to show both flightcrew and ground crew personnel the wind direction and velocity at any given time. Additionally, the location of the ground station should be such that it is free of rotor downwash at all times. Motion picture or phototheodolite and radio equipment will be necessary to properly conduct the test program. The use of telemetry equipment is desirable if the location of the test site and the magnitude of the test program make it practical. (ii) Airborne Equipment (Test Rotorcraft). Necessary installed test equipment may include photopanels and/or recorders for recording engine parameters, control positions, landing gear loads, landing gear deflections, airspeed, altitude, and other variables. An external light attached to the rotorcraft (or any other means of identifying the engine failure point to the ground camera or phototheodolite) is needed to identify the exact time of engine failure and may also be used to synchronize the ground recorder with the airborne recorded data. (2) Analytical Prediction . The HV diagram can be estimated by analytical means and this is recommended prior to test. HV, however, is the least predictable of all rotorcraft performance and because of this, the expansion and extrapolation of test data must be done with great care. Test weight may not be extrapolated. All test points should be approached conservatively with some speed or altitude margin. If the applicant has conducted a comprehensive HV flight test program to validate his analytical predictions, much preliminary testing can be eliminated. In any case, the maximum allowable extrapolation from flight test conditions is 2,000 feet density altitude, and an approved analytical and/or simulation method must be utilized for extrapolation. (3) Power. (i) The appropriate power level before engine failure for the low and high hover points is simply the power required to hover at the prevailing hover conditions. The appropriate power condition prior to failure of the engine for points below the knee is takeoff power or a lower value if approved as an operating procedure. For cruise or “fly-in” points above the knee, the appropriate condition is power required for level flight. (ii) The applicable power failure conditions are listed in § 29.79(b). Power should be completely cut for normal category rotorcraft. For multiengine rotorcraft with Category A engine isolation, only one engine need be failed and the desired topping power (for the remaining engine(s)) should be set prior to the test. This power value will need adjustment as ambient conditions change. The power can be takeoff power (TOP), 2 ½-minute power, or some calculated lower power for simulating hot day or higher density altitude conditions. Power is verified and recorded by the pilot by “topping” the engine(s) prior to engine failure tests. Care must be taken to ensure Page B - 39 ### Page 77 that this power value is no more than that which would be delivered by a minimum specification engine under the ambient conditions to be approved. (4) Test Loadings . Weight extrapolation is not permitted for HV. Therefore, the test weight must be closely controlled. Ballast or fuel should be added frequently to maintain the weight within -1 to +5 percent when testing final points. Ordinarily, tests are conducted at a mid center of gravity unless a particular loading is expected to be particularly critical. (5) Landing Gear Loads. (i) Instrumented landing gear can be a great help in evaluating test results. This information can be telemetered to a ground station or otherwise recorded and displayed for direct reference following each landing. (ii) Any landing which results in permanent deformation of aircraft structure or landing gear beyond allowable maintenance limits is considered an unsatisfactory test point. (6) Piloting Considerations . In verifying the HV diagram, the minimum certificated instrument panel display and minimum crew should be used in order not to mislead the operational pilot who has no test equipment available and may have no copilot to assist. Three distinctly different flight profiles are utilized in developing the diagram. (i) High Hover. A stabilized out-of-ground-effect (OGE) hover condition prior to power failure is essential. A minimum 1-second time delay between power failure and initial control actuation is utilized. Following the time delay, the primary concern is to quickly lower collective and to gain sufficient airspeed to allow an effective flare approaching touchdown. While the immediate development of airspeed is necessary, the dive angle must be reasonable and must be representative of that expected in service. While initial aircraft attitude will vary between models and with changing conditions, 10°-20° has been previously applied as a maximum allowable nose down pitch attitude. Use of greater attitudes could result in a diagram which is difficult to achieve and unrealistic for operations in service. Initial testing should start relatively high with gradual lowering of height to the final high hover altitude. A stabilized OGE hover condition prior to power failure is essential. If a stabilized high hover condition cannot be achieved prior to the engine cut, then this point should be tested from a minimum level flight speed. This will result in an open-ended HV diagram. A smoke source or balloon on a long cord is highly desirable since the wind can vary significantly from surface observations to typical high hover altitudes. Vertical speed must be very near zero at the throttle chop. Any climb or sink rate can have a significant influence on the success of the test point. Use of a radar altimeter with a cross check to barometric altitude is essential. Page B - 40 ### Page 78 (ii) Low Hover . From the low hover position there is no flare capability and little time for collective reaction. No time delay is applied other than normal pilot reaction. For typical designs the collective may not be lowered after power failure. Lowering of the collective is not permitted because it is not a pilot action which could be expected if an engine failed without notice during a hovering condition in service. Initial lowering of collective immediately after power failure can result in a very high, unconservative low hover height that is unrealistic for operational conditions. If, however, a design is such that a 1-second pilot delay after power failure could be achieved without any appreciable descent, a slight lowering of collective could be allowed. (iii) Takeoff Corridor . Normal pilot reaction is applied when the engine is made inoperative. At low speeds, collective may be lowered quickly to retain RPM and minimize the time between power failure and ground contact. If airspeed is sufficient for an effective flare, the aircraft is flared to reduce airspeed, retain rotor RPM, and control vertical speed prior to touchdown. Considerable surface area may be needed for a sliding or rolling stop. (iv) Additional Considerations . The “in-between” points utilize similar techniques. The cruise or “fly-in” points are similar to the high hover point although the steep initial pitch attitudes are not needed as altitude is decreased and airspeed is increased along the curve. The low speed points along the takeoff corridor are similar to the low hover point except that the collective may be quickly lowered and some flare capability may be used as the “knee” is approached. The pilot should be proficient in all normal autorotation landings before conducting HV tests in a single-engine rotorcraft. (7) Ground Support . Motion picture or theodolite coverage and ground safety equipment are necessary. Communication capability among these elements should be provided. Use of a phototheodolite to compare height/speed with cockpit observations is very desirable. (8) Verifying the HV Diagram. (i) A sufficient number of test points must be flown to verify the diagram. The key areas are the knee, high altitude hover, low altitude hover, and low altitude high speed flight. Test points with excessive gear loads, exceptional skill requirements, winds above permissible levels, rotor droop below approved minimum transient RPM, damage to the rotorcraft, excessive power, incorrect time delay, etc., cannot be accepted. (ii) After the HV diagram is def ined, it should be ascertained that the corridor permits takeoffs within ±5 knots of the recommended takeoff profile. (9) Flight Manual . The flight manual should list any procedures which may apply to specific points (e.g., high speed points) and test conditions, such as runway surface, wave height for amphibious tests, marginal areas of controllability or landing Page B - 41 ### Page 79 gear response, etc. The HV curve should be presented in the RFM using actual altitude above ground level and indicated airspeed. (10) Night Evaluation . If a rotorcraft is to be certified for night operation, a night evaluation is required. Simulated engine failures should be conducted along the recommended takeoff path. Landings should also be qualitatively evaluated with an engine failed. Engine failures at critical HV conditions are not required. The intent is to show adequate visibility using aircraft and/or runway lights without requiring a duplication of the daytime HV test program. (11) Water Landings . For amphibious float-equipped rotorcraft, day and night water landings should be conducted under critical loading conditions with an engine failed. Engine failures should be conducted along the recommended takeoff path. Engine failures at critical HV conditions are not required. The intent is to show similarity to test results over land without requiring a duplication of the HV test program. Page B - 42 ### Page 80 (OCR) HEIGHT LEVEL FLIGHT (H} -KNEE ЛАКЕОНЕ **TAKEOFF CORRIDOR** **HIGH SPEED I** SPEED (V) FIGURE AC 27.79-1 HEIGHT - VELOCITY DIAGRAM Page B - 43 ### Page 81 (OCR) Approx Constant GW/e Height Velocity Weight —A— CONSTANT HV DIAGRAM, VARIABLE WEIGHT Increasing Density Altitude Height Velocity — B CONSTANT WEIGHT Increasing Gross Weight Height Velocity CONSTANT DENSITY ALTITUDE FIGURE AC 27.79-2 ALTITUDEMEIGHT ACCOUNTABILITY Page B - 44 ### Page 82 AC 27.79A. § 27.79 (Amendment 27-21) LIMITING HEIGHT-SPEED ENVELOPE. a. Explanation . Amendment 27-21 to the regulation redefines the required weight for establishing the HV envelope at altitudes above sea level. b. Procedures . All of the policy material pertaining to this section remains in effect. In addition, the following applies: (1) The rotorcraft height-velocity envelope should be established for the maximum gross weight at sea level. At altitudes above sea level, the envelope should be established at not less than the maximum operating weight or OGE hover weight, whichever is lower. If a weight below the OGE hover weight is selected, by definition, that selected weight becomes the maximum operating weight for the rotorcraft at that altitude. (2) If the HV envelope is established for a maximum altitude less than 7,000 feet, by definition, the maximum takeoff and landing altitude for the rotorcraft may be no higher than that maximum HV altitude. Hover performance information should not be presented for altitudes above the maximum altitude for which the HV envelope is established. Page B - 45 ### Page 83 **SUBPART B - FLIGHT** **FLIGHT CHARACTERISTICS** AC 27.141. § 27.141 (Amendment 27-21) FLIGHT CHARACTERISTICS--GENERAL. a. Explanation . (1) This section prescribes the general flight characteristics required for certification of a normal category rotorcraft. Specifically, it states that the rotorcraft shall comply with the flight characteristics requirements at all approved operating altitudes, gross weights, center of gravity locations, airspeeds, power, and rotor speed conditions for which certification is requested. The reference to altitude in § 27.141(a)(1) refers to density altitude. Density altitude is, of course, a function of pressure altitude and ambient temperature, hence the need to account for ambient temperature effects. Additional flight characteristics required for instrument flight are contained in AC 27 Appendix B of this advisory circular. (2) Generally, the aircraft structural (load level) survey accounts for takeoff power values at speeds up to and including VY. At speeds above VY, maximum continuous power is assumed. Stress to rotating components usually increases with airspeed and power. If the takeoff power rating exceeds the maximum continuous power rating, and the structural survey has been conducted under the assumption that takeoff power is not used at speeds above VY, the Rotorcraft Flight Manual must limit takeoff power to speeds of VY and below. If takeoff power is structurally substantiated throughout the flight envelope, and appropriate portions of the controllability, maneuverability, and trim requirements of §§ 27.141 through 27.161 are met at takeoff power levels, no flight manual entry is needed. Obviously if transmission limits for maximum continuous (MC) and takeoff power coincide, no special action is needed. (3) During the flight characteristics testing, the controls must be rigged in accordance with the approved rigging instructions and tolerances. The control system rigging must be known prior to testing. In addition to the normal rigging procedures, any programmed control surfaces, which may be operated by dynamic pressure, electronics, etc., must also be calibrated. During the flight test program, it is frequently necessary to rig a control, such as the swashplate or tail rotor blade angle, to the allowable critical extreme of the tolerance band. For example, it would be necessary to rig the tail rotor to the minimum allowable blade angle if meeting the requirements of §27.143(c) would be in question. The same consideration must be given to all rotorcraft controls and movable aerodynamic surfaces where questionable compliance with the regulations may exist. If the rotor-induced vibration characteristics of the rotorcraft are significantly affected and require time-consuming rigging for such things as acceptable ride comfort, then the rotor(s) should be rigged to the allowable extreme tolerance limits to determine compliance, for example, with § 27.251. Page B - 46 ### Page 84 (4) During the FAA/AUTHORITY fli ght test program, the crew should be especially alert for conditions requiring great attentiveness, high skill levels, or exceptional strength. If any of these features appear marginal, it is advisable to obtain another pilot’s opinion and to carefully document the results of these evaluations. Section 27.141(b) provides the regulatory basis for these strength and skill requirements. The general requirements for a smooth transition capability between appropriate flight conditions are also included in § 27.141(b). These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft. (5) For night or IFR approval, § 27.141( c) contains the general regulatory reference, which requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this advisory circular. b. Procedures . none AC 27.143. § 27.143 (Amendment 27-21) CONTROLLABILITY AND MANEUVERABILITY . a. Explanation. (1) This regulation contains the bas ic controllability requirements for normal category rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for control and for maneuverability are summarized in § 27.143(a) which is largely self-explanatory. The hover condition is not specifically addressed in § 27.143(a)(2) so that the general requirement may remain applicable to all rotorcraft types, including those without hover capability. For rotorcraft, the hover condition clearly applies under “any maneuver appropriate to the type.” (2) Paragraphs (b) through (e), § 27.143, include more specific flight conditions and highlight the typical areas of concern during a flight test program. (i) Section 27.143(b) specifies flight at V NE with critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at VNE for nosedown pitching of the rotorcraft and for adequate roll control. Nosedown pitching capability is needed for control of gust response and to allow necessary flight path changes in a nosedown direction. Roll control is needed for gust response and for normal maneuvering of the aircraft. In the past, 10 percent control travel margin has been applied as an appropriate minimum control standard. The required amount of control power, however, has very little to do with any fixed percentage of remaining control travel. There are foreseeable designs for which 5 percent remaining is adequate and others for which 20 percent may not be enough. The key is, can the remaining longitudinal control travel at VNE generate a clearly positive nosedown pitching moment, and will the remaining lateral travel allow at least 30° banked turns at reasonable roll rates? Moderate lateral control reversals should be included in this evaluation and since available roll control can diminish with sideslip, reasonable out of trim conditions Page B - 47 ### Page 85 (directionally) should be investigated. This “control remaining” philosophy must also be applied for other flight conditions specified in this section. (ii) Section 27.143(c) requires a minimum control capability for hover and takeoff in winds of 17 knots from any azimuth. Control capability in wind from zero to at least 17 knots must also be shown for any other appropriate maneuver near the ground such as rolling takeoffs for wheeled rotorcraft. These requirements must be met from standard sea level conditions to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less. On rotorcraft incorporating a tail rotor, efficiency of the tail rotor decreases with altitude so that a given sideward flight condition requires more pedal deflection, a higher tail rotor blade angle, and more horsepower. Hence, directional capability in sideward flight (or at critical wind azimuth) is most critical during testing at a high altitude site. (iii) Section 27.143(d) requi res adequate controllability when an engine fails. This requirement specifies conditions under which engine failure testing must be conducted and includes minimum required delay times. (A) For rotorcraft which meet the engine isolation requirements of transport Category A, demonstration of sudden complete single-engine failure is required at critical conditions throughout the flight envelope including hover, takeoff, climb at VY, and high speed flight up to VNE. Entry conditions for the first engine failure are engine or transmission limiting maximum continuous power (or takeoff power where appropriate) including reasonable engine torque splits. For multiengine Category A installations (three or more engines) subsequent engine failures should be conducted utilizing the same criteria as that used for first-engine failure. The applicant may limit his flight envelope for subsequent failures. Initial or sequential engine failure tests are ordinarily much less severe than the “last” engine failure test required by § 27.75(b). The conditions for last-engine failure are maximum continuous power, or 30-minute power if that rating is approved, level flight, and sudden engine failure with the same pilot delay of 1 second or normal pilot reaction time, whichever is greater. (B) For rotorcraft without transport Category A engine isolation, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to VNE (power-on) and conditions of hover, takeoff, and climb at VY. Maximum continuous power is specified prior to the failure for the cruise condition. Power levels appropriate to the maneuver should be used for other conditions. The corrective action time delay for the cruise failure should be 1-second or normal pilot reaction time (whichever is greater). Cyclic and directional control motions which are apart of the pilot task of flight path control are normally not subject to the 1-second restriction; however, the delay is always applied to the collective control for the cruise failure. If the aircraft flying qualities and cyclic trim configuration would encourage routine release of the cyclic control to complete other cockpit tasks during cruise flight, consideration should be given to also holding cyclic fixed for the 1-second delay. Although the same philosophy could be extended to the directional controls, the likelihood of the pilot having his feet away from the pedals is Page B - 48 ### Page 86 much lower, unless the aircraft has a heading hold feature. Rotor speed at execution of the cruise condition power failure should be the minimum power-on value. The term “cruise” also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies maximum continuous (MC) power, it does not limit engine failure testing to MC power. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions. Following power failure, rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits. (iv) Section 27.143(e) addresses the special case in which a V NE (power-off)is established at an airspeed value less than VNE (power-on). For this case, engine failure tests are still required at speeds up to and including VNE (power-on), and the rotorcraft must be capable of being slowed to VNE (power-off) in a controlled manner with normal pilot reactions and skill. There is, however, no controllability requirement for stabilized power-off flight at speeds above 1.1 VNE (power-off) when VNE (power-off) is established per § 27.1505(c). (v) Application of the controllabilit y requirement for pitch, roll, and yaw at speeds of 1.1 VNE (power-off) and below is similar to that described above for power-on testing at VNE. Sufficient directional control must exist to allow straight flight in autorotation during all approved maneuvers including 30° banked turns up to VNE (power-off) with some small additional allowance for gust control. Adequate controllability margins must exist in all axes throughout the approved autorotative flight envelope. Testing to VNE at MC power per § 27.143(b), 1.1 VNE at power for 0.9 VH per § 27.175(b) or § 27.1505, and to 1.1 VNE (power-off) in autorotation per § 27.143(e) should be sufficient to assure adequate control margin during a descent condition at high speed and low power. The high speed, power-on descent condition should be checked for adequate control margin as a “maneuver appropriate to the type.” There has been one instance where insufficient directional pedal was available to maintain a reasonable trimmed sideslip angle with low power at very high speeds, and a case where there was insufficient forward and lateral cyclic available to reach the power on VNE. The insufficient directional pedal margin was due to the offset vertical stabilizers. The lack of cyclic stick margin was because the cyclic stick migrated to the right as power was reduced, and the control limits were circular. This provided less total available forward cyclic stick travel when the cyclic was moved right and forward about 45° from the center position. Each of the above rotorcraft was certificated with a rate of descent limitation to preclude operation in the control-limited area. (vi) An evaluation of the emergency descent capability of the rotorcraft should be made, either analytically or through flight test. Areas of consideration are the rate of descent available, the maximum approved altitude, and the time before a catastrophic failure following the loss of transmission oil pressure or other similar failure. Each rotorcraft should have the capability to descend to sea level and land from the maximum certificated altitude within the time period established as safe following a critical failure. If the time period does not permit a sea level landing, the maximum Page B - 49 ### Page 87 height above the terrain must be specified in the limitation section of the Rotorcraft Flight Manual. (3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 27.672). Examples include stability augmentation systems (SAS), stability and control augmentation systems (SCAS), automatic flight control systems (AFCS), devices to provide or improve longitudinal static stability such as a pitch bias actuator (PBA), yaw dampers, and fly-by-wire elevator or stabilator surfaces. These systems all use actuators of some type, and are subject to actuator softover and hardover malfunctions. The flight control system should be evaluated to determine whether an actuator jammed in an extreme position would result in reduced control margins. Generally, if the flight control system stops are between the actuator and the cockpit control, the control margin will be affected. If the control stops are between the actuator and the rotor head, the control margins may not be affected, but the location of the cockpit control may be shifted. This could produce interference with other items in the cockpit. An example of this would be a lateral actuator jammed hardover causing a leftward shift in the cyclic stick position. Interference between the cyclic stick, the pilot’s leg, and the collective pitch control could reduce the left lateral control available and reduce left sideward flight capability. In the case of fly-by-wire surfaces, both the high speed forward flight controllability and the rearward flight capabilities could be affected. Flight control systems that incorporate automatic devices should be thoroughly evaluated for critical areas. Every failure condition that is questionable should be flight tested with the appropriate actuator fixed in the critical failure position. These failures may require limitations of the flight envelope. Any procedure or limitation that must be observed to compensate for an actuator hardover and/or softover malfunction should be included in the Rotorcraft Flight Manual. b. Procedures . (1) Flight test instrumentation shoul d include ambient parameters, all flight control positions, rotor RPM, main and tail rotor flapping (if appropriate), engine power instruments, and throttle position. Flight controls that are projected to be near their limits of authority should be rigged to the most adverse production tolerance. A very accurate weight and balance computation is needed along with a precise knowledge of the aircraft’s weight/CG variation as fuel is burned. (2) The critical condition for V NE controllability testing is ordinarily aft CG, MC power, and minimum power-on rotor RPM, although power and RPM variations should be specifically evaluated to verify their effects. The turbine engine is sensitive to ambient temperatures which affect the engine’s ability to produce rated maximum continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit maximum continuous power would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward CG should be spot checked for any “tuck under” tendency at high speed. The VNE controllability test is normally Page B - 50 ### Page 88 accomplished shortly after the 1.1 VNE (or 1.1VH) point obtained during stability tests required by § 27.175(b). Controllability must be satisfactory for both conditions. If VNE varies with altitude or temperature, VNE for existing ambient conditions is utilized for the test. Extremes of the altitude/temperature envelope should be analyzed and investigated by flight test. (3) The critical condition for controllabilit y testing in a hover is ordinarily forward CG at maximum weight with minimum power-on rotor RPM. For rearward flight testing of configurations where the forward CG limit varies with weight, low or high gross weight may be critical. Lateral CG limits should also be investigated. A calibrated pace vehicle is needed to assure stabilized flight conditions. Surface winds should be less than 3 knots throughout the test sequence. Testing can be done in higher stabilized wind conditions (gusting less than 3 knots); however, these conditions are very difficult to find and the method is very time consuming due to the necessity of waiting for stabilized winds. Testing in calm winds is preferred. Hover controllability testing should be accomplished with the lowest portion of the rotorcraft at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. Although the necessary yaw response will vary somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the rotorcraft is held in the most critical position relative to wind. (4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of his engine, drive system, and rotor limits. These limits were established during previous ground and flight tests and should be specified in the TIA. Particular attention should be given to minimum stabilized and minimum transient rotor RPM limits. These values must be included in the TIA and should be approached gradually with a build-up in time delay unless the company testing has completely validated all pertinent aspects of engine failure testing. On Category A installations, the maximum power output of each engine must be limited so that when an engine fails and the remaining engine(s) assume the additional load, the remaining engine(s) are not damaged by excessive power extraction and over-temping. This is needed for compliance with § 27.903(b). The propulsion engineer should have assured that this feature was properly addressed in the engine and drive system substantiation; however, it must be assumed that for some period of time the pilot may extract maximum available power from the remaining engine(s) when an engine fails during critical flight maneuvers. Substantiation of this feature should be accomplished primarily by engine and drive system ground tests. (5) Longitudinal cyclic authority at V NE with any power setting must permit suitable nosedown pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft’s response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft Page B - 51 ### Page 89 motion, the results are unacceptable; e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral cyclic to correct roll in conjunction with forward cyclic to correct pitchup is satisfactory. Obviously during the conduct of these tests, all available techniques should be utilized when the pilot finds himself “out of control.” However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion. (6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all-engine-inoperative requirements of § 27.75. AC 27.151. § 27.151 (Amendment 27-21) FLIGHT CONTROLS . a. Explanation . Excessive breakout or preload in the flight controls produces control system force discontinuities which result in increased workload and controllability problems for the pilot. Similarly, excessive freeplay results in lost motion which increases pilot workload and, in an extreme case, could lead to a hazardous pilot-induced oscillation. In some designs friction can provide a positive contribution to the function of the flight controls (e.g., masking aerodynamic feedback in reversible systems). At some point, friction will have a detrimental effect on the pilot’s ability to properly control the machine. In the case of an irreversible design equipped with an artificial force feel system in pitch and roll, excessive friction can mask a shallow force gradient making positive stick centering and control force static stability difficult if not impossible to demonstrate. In such an instance, the initial choice of fixes might include implementation of a steeper force gradient or addition of a force preload. Unfortunately, these solutions often lead to the kinds of problems discussed earlier. Care must therefore be exercised during the initial design phase to ensure that the components and characteristics of the flight control system are well matched. b. Procedures . Regardless of the flight control system sophistication, it is important that the test pilot understand the system configuration prior to flight evaluation. Appropriate mechanical characteristics should be documented. For VFR aircraft, the mechanical characteristics are typically assessed in flight on a qualitative basis. If a controllability or workload problem is identified, a more detailed investigation would be necessary. Since IFR certification rules include specific trim and force requirements, a more quantitative investigation of mechanical characteristics is normally conducted. The constantly varying feedback forces of reversible flight control systems generally make such designs unsuitable for IFR application. Irreversible system mechanical characteristics can often be partially documented on the ground with external hydraulic and electrical power supplies connected to the aircraft. Knowledge of the breakout, friction, and force gradient characteristics prior to flight can be useful to the pilot during flight evaluation of the system. Page B - 52 ### Page 90 AC 27.161. § 27.161 (Amendment 27-21) TRIM CONTROL . a. Explanation . (1) The pilot has many tasks to per form with each hand during sustained flight conditions. The trim requirement is intended to reduce the physical demands to maintain a given flight condition. It is not intended to require that control forces be reduced to zero by the trim control during dynamic maneuvers such as takeoff acceleration. (2) A number of devices may be used to produce the necessary trim characteristics. One popular method of meeting this requirement is through the use of control balance springs in conjunction with a small amount of built-in control system friction. Other methods include use of friction, magnetic brakes, bungees, and irreversible mechanical schemes. (3) This regulation is not intended to r equire zero friction or zero breakout force in the control system, nor is it intended to require automatic control recentering. The regulation, in fact, specifically prohibits excessive high friction or high breakout forces which would produce undesirable discontinuities in the primary control force gradient. b. Procedures . (1) If comprehensive company flight te st data are available, compliance with this requirement can quickly be found by spot checking extreme center of gravity loadings. Trim tests can ordinarily be done during the course of other flight test activities. To conduct the test, briefly release the control at the required flight conditions and determine that the control does not move. The words “any appropriate speed” ordinarily include any speed from hover to VH. If the control system trim device might be subject to temperature or humidity effects, these should be investigated at a minimum of two altitude extremes and during several test phases. (2) If a pilot controllable variable fric tion device is incorporated, compliance with this requirement must be shown at the minimum adjustable value. The maximum value of adjustable friction should not completely lock the flight controls. (3) Continued compliance with this requirement should be ensured through a production procedure. If minimum friction or centering springs are used, it is desirable for the manufacturer to include some adjustment capability for production differences. The explanation and procedures discussed here are applicable for VFR approval under § 27.161. For additional IFR trim requirements, refer to AC 27 Appendix B. AC 27.171. § 27.171 STABILITY: GENERAL . a. Explanation . This section is intended to require a manageable pilot workload for the minimum crew under foreseeable operating conditions. Page B - 53 ### Page 91 b. Procedures . (1) Compliance with the requirements of this section can often be obtained for the VFR condition without any specific or designated flight testing. If the rotorcraft is marginal in regard to pilot strain and fatigue, the FAA/AUTHORITY pilot should be assured, through special tests if necessary, that the aircraft can be satisfactorily flown throughout the maximum endurance capabilities of the rotorcraft including night and turbulence conditions if those are critical. This test should be conducted with minimum required systems in the aircraft and with minimum flightcrew. (2) Reasonable failure conditions wh ich add to pilot workload, strain, and fatigue should be evaluated (electrical, hydraulic, and mechanical failures, etc.). The necessary times associated with flight with a failed system must be appropriate to the flight manual procedures for each failure. A failure condition requiring immediate landing would obviously require shorter evaluation time than a condition allowing continued flight to destination. (3) IFR approvals necessitate a care ful evaluation of paragraphs b (1) and (2) above. In IFR operations, weather conditions frequently necessitate continued flight to destination or diversion to alternate airports with critical failures. Immediate landing may not be feasible. The evaluating pilot must ensure pilot strain and fatigue are acceptable during typical flight profiles for each type of operation to be approved. AC 27.173. § 27.173 (Amendment 27-21) STATIC LONGITUDINAL STABILITY. a. Explanation . (1) This rule contains control r equirements for both stability and control. Paragraph (a) contains the basic control philosophy necessary for all civil aircraft. Forward motion of the cyclic control must produce increasing speeds, and aft motion must result in decreasing speeds. For rotorcraft, this is accomplished with throttle and collective held constant. Rotorcraft with either highly stable or highly unstable static longitudinal stability characteristics can typically comply with the basic requirement for control sense of motion. However, the intent and interpretation of this paragraph is to provide a stable stick position versus airspeed gradient. Therefore, a stabilized airspeed less than the trim speed requires a cyclic stick position aft of the trim stick position, and a stabilized airspeed greater than the trim speed requires a cyclic stick position forward of the trim speed stick position. (2) The remainder of § 27.173, thr ough reference to § 27.175, contains the basic control position requirements necessary to establish a minimum level of static longitudinal stability. Positive stability is found for conditions of climb, cruise, and autorotation in § 27.175 by requiring a stable stick position gradient through a specified speed range. A defined level of instability is permitted for the hovering condition. Page B - 54 ### Page 92 b. Procedures . (1) The control requirement of this se ction is so essential to basic flight mechanics that compliance may be found during conventional flight testing for compliance with other portions of the regulations. No special or designated testing should be required. (2) The procedures necessary to ensure compliance with the stability requirements of this section are contained under § 27.175. AC 27.175. § 27.175 (Amendment 27-21) DEMONSTRATION OF STATIC LONGITUDINAL STABILITY . a. Explanation . (1) This rule incorporates the specific flight requirements for demonstration of static longitudinal stability. Specific loadings, configurations, power levels, and speed ranges are stated for conditions of climb, cruise, autorotation, and hover. (2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in § 27.173 which states that throttle and collective pitch must be held constant is appropriate for administration of this rule, and the specified power in § 27.175(a), (b), and (c) should be considered as power established at initial trim conditions. This will result in slightly higher or lower torque readings at “off trim” conditions. Collective and throttle controls are held constant when obtaining data during climb, cruise, and autorotation tests. (3) The effects of rotor RPM on autorotative static stability should be determined and positive stability demonstrated for the most critical RPM. Values for RPM can be expected to change as airspeed is varied from the “trimmed” condition. The manufacturer’s recommended autorotation airspeed is ordinarily used for trim. (4) Hovering is considered a flight maneuver for which the pilot repeatedly adjusts collective to maintain an approximately constant altitude above the ground. For hover stability tests, collective and throttle adjustments are made as necessary to maintain an approximately constant height above the ground. Also, a limited amount of negative longitudinal control travel is allowed with changes in speed. b. Procedures . (1) Instrumentation. (i) Sensitive control position inst rumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when Page B - 55 ### Page 93 using a “before and after” method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The figure AC 27.175-1 plot is a typical presentation of longitudinal static stability. (ii) Other necessary parameter s include pressure altitude, ambient temperature, and indicated airspeed (pace vehicle or theodolite speed for hover tests). For hover tests, hover height (radar altitude if available) and surface winds should be documented. Two-way communication with a pace vehicle is highly desirable. Ground safety equipment is desirable. (2) Ambient Conditions. Smooth air is necessary for stability testing. Allowable wind conditions for hover stability testing are the same as those for hover controllability tests. Extrapolation is covered in paragraph AC 27.45. (3) Loading. Aft center of gravity (CG) is ordinarily critical for longitudinal stability testing, although high speed flight and hover should be checked at full forward CG and maximum weight. At aft CG, light or heavy weight conditions can be critical. The manufacturer’s flight data should be reviewed to determine critical loading conditions. (4) Conducting the Test. (i) The rotorcraft should be established in the desired configuration and flight condition (climb, cruise, autorotation) with the required power and rotor speed at the trim airspeed. The collective stick should be fixed in that position, usually by applying sufficient friction to ensure that it is not inadvertently moved. For autorotative tests, a rotor speed should be selected so that the variations in rotor speed as airspeed and altitude change do not exceed the allowable limits. This point is recorded as the trim point. Airspeed is then increased or decreased in about 10-knot increments, stabilizing on each speed and recording the data. At least two points on each side of the trim speed should be taken. (ii) The cruise test should be accomplished by first determining V H (level flight speed at maximum continuous power) at the test altitude. Then reduce power to establish a level flight trimmed condition at 0.9 VH (or 0.9 VNE if lower). This point is then recorded as the trim point. The collective pitch and throttle must remain fixed at the trim setting for the remainder of the test. The airspeed is then varied above and below the trim speed using the cyclic control to climb or dive slightly. (iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually ±2,000 feet), first increasing airspeed as data points are collected, then decreasing speed through the same altitude band. It will probably not be Page B - 56 ### Page 94 possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit. Generally, it will be possible to acquire all the high speed points on one pass and the low speed points on the second. (iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise, data can be obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise data because collective and throttle are not moved as airspeed is changed at a constant altitude. A typical sequence of speeds that could produce these results would be: 150 (VH), 135 (0.9VH) trim speed, 125, 145, 115, 155, 105, and 165. (v) For rotorcraft with high rates of climb, a series of climbs, each at a different speed, may be required through a given altitude, utilizing sensitive instrumentation to ensure collective position is the same for each data point. In autorotation, a similar case arises and a series of descents, each at a different speed, may be required through a given altitude band, using sensitive instrumentation to ensure a repeatable collective position. (vi) Hover tests should be conducted by maintaining an approximately constant altitude above the ground at the hover height established for performance purposes. The test altitude above the ground may be increased to provide reasonable ground clearance during rearward flight. Groundspeed is varied using a pace vehicle, theodolite, or other velocity measuring equipment. A pace vehicle is an aid in maintaining an accurate hover height. The pilot can accurately maintain height by controlling his sight picture of the pace vehicle (level with the roof, antenna, etc.). Hover stability tests are ordinarily conducted in conjunction with hover controllability tests because instrumentation and facilities are essentially the same. (vii) Normally, climb, cruise, and aut orotation tests should be conducted at low, medium, and high altitudes. See paragraph AC 27.45 for guidance on interpolation and extrapolation. High speed stability has been critical during cold weather testing. In two recent models, VNE at cold temperatures has been limited by the stability requirements of § 27.175(b). Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather VNE to tip Mach number values demonstrated during warm weather testing. (viii) Hover stability should be verifi ed at low altitude and, if required, at high altitude. Refer to paragraph AC 27.45b(2) for guidance on expansion and extrapolation of altitude. Page B - 57 ### Page 95 (OCR) o CRUISE • CLIMB **A AUTOROTATION** **O HOVER** SHADED POINTS REPRESENT TRIM 100 LONGITUDINAL CYCLIC POSITION (% FROM FULL AFT) 20 000000. 000600 10 20 40 60 80 100 120 _ 140 160 CALIBRATED AIRSPEED (KNOTS) FIGURE AC 27.175-1 STATIC LONGITUDINAL STABILITY Page B - 58 ### Page 96 AC 27.177. § 27.177 (Amendment 27-21) STATIC DIRECTIONAL STABILITY. a. Explanation . This rule requires that positive static directional stability be demonstrated at the trim airspeeds defined in § 27.175. The trim speed for climb is VY and for cruise 0.9VH or 0.9VNE (whichever is less). b. Procedures . (1) Tests for static directional stabilit y require instrumentation for pedal position and sideslip angle. To obtain accurate sideslip angle and airspeed information, a “yaw boom” is usually installed for the purpose of mounting a sideslip vane and swiveling airspeed pilot head outside the main rotor downwash region of influence. Special care should be taken to ensure that the yaw boom installation has been verified to be structurally adequate and free of dynamic instabilities for all combinations of airspeed and rotor speed likely to be experienced during the static directional evaluation. For some installations, the instrumentation yaw boom may influence the flying qualities of the rotorcraft itself. Thus, it is advisable to correlate yaw string displacement or slip indicator ball widths of skid with yaw boom sideslip angle, and then repeat a few critical points with the yaw boom removed. (2) For some rotor system designs, the main and tail rotor flapping angle may be a critical instrumentation requirement for static directional testing. Both main and tail rotor flapping may increase dramatically at high airspeeds with increasing sideslip angle. Therefore, for rotor systems exhibiting this characteristic, flapping should be monitored carefully during the sideslip maneuver to avoid exceeding limitations. Static directional stability is normally defined in terms of pedal displacement required to maintain a steady heading sideslip. A single-rotor rotorcraft flying in coordinated flight will exhibit a small inherent sideslip due to tail rotor thrust and fuselage/main rotor sideforces. This condition is normally taken as trim with the inherent sideslip angle noted. Airspeeds should be the trim values described above. The procedures used to establish and maintain the steady heading sideslip can significantly influence test results. A generally accepted technique follows: (i) Stabilize at the trim point, and note indicated airspeed. (ii) Record trim conditions incl uding inherent sideslip. Maintain fixed collective and throttle for the remainder of the maneuver. (iii) Smoothly apply directional control and coordinate with lateral control to establish the desired sideslip angle. A steady track can best be ensured by maintaining a track over a straight landmark on the ground such as a section line or straight segment of powerline or highway. (iv) Because drag increases with si deslip, the aircraft will decelerate. The trim airspeed should be maintained by entering a slight dive or decreased rate of climb. If a boom airspeed system is not used, a “standard” airspeed system will become Page B - 59 ### Page 97 excessively inaccurate at about 10° of sideslip. Rapidly yawing to the desired sideslip angle and noting and maintaining the new “standard” indicated airspeed may give adequate data. Control positions (directional as a minimum), sideslip angle, rotorspeed, airspeed, rate of descent, amount of ball deflection, and bank angle should be recorded. The pilot should note the physical sideforce feel experienced. The rule requires that sufficient cues accompany sideslip to alert the pilot when approaching sideslip limits. A minimum of two sideslip data points on each side of the trim point should be obtained to adequately define the slope of the pedal displacement versus sideslip angle relationship. (v) Static directional stability plot s can be expected to differ slightly on either side of the inherent sideslip angle. Positive static directional stability is indicated by increased left pedal displacement for a larger right sideslip and, conversely, increased right pedal for a larger left sideslip angle. Page B - 60 ### Page 98 **SUBPART B - FLIGHT** **GROUND AND WATER HANDLING CHARACTERISTICS** AC 27.231. § 27.231 GROUND AND WATER HANDLING CHARACTERISTICS-- GENERAL. a. Explanation . The rule states: “The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.” In addition, §§ 27.235, 27.239, and 27.241 contain specific requirements concerning ground and water handling characteristic evaluations. b. Procedures . (1) During the flight test program and the F&R program (§ 21.35(b)(2)), the rotorcraft will be subjected to evaluations at various weight and CG conditions. Any uncontrollable tendencies found during these test programs must be corrected. (2) Controllable or damped vibrations or oscillations on the ground or in the water are acceptable, provided the design limits of the rotorcraft are not exceeded. (3) Any significant vibration or o scillation characteristics found during tests should be described in the test report, and the rotorcraft flight manual should contain appropriate descriptions and procedures to describe and either avoid or handle significant characteristics. (4) For rotorcraft equipped with wheel gear, the evaluation should include takeoff, landing, and taxi at the maximum speed and at CG extremes. If a nose or tail wheel lock/swivel control is installed, each position should be evaluated for limiting takeoff, landing, and taxi speeds. Maximum substantiated speed values should be included in the RFM as limitations. (5) For water operations, the wave hei ght and frequency or “sea state” should be included as a limitation or, if no limit was reached during testing, the demonstrated values should be placed in the Performance Section of the RFM. Information or limits on the allowable “sea state” for rotor startup and shutdown should also be included. AC 27.235. § 27.235 TAXIING CONDITION . a. Explanation . The rotorcraft is designed for certain landing load factors (§§ 27.471 and 27.473). The rotorcraft must not attain a load factor in excess of the design load factor when taxied over the roughest ground that may reasonably be expected in normal operation at the expected taxi speeds. This rule applies to wheel landing gear equipped rotorcraft. Page B - 61 ### Page 99 b. Procedures . The structural substantiation data contain the allowable design limits for the rotorcraft. A calibrated accelerometer or load factor “g” meter should be installed as near as practicable to the rotorcraft CG to record the maximum vertical load factor attained. Instrumentation of the landing gear and/or related structure may also be an acceptable means of showing compliance. (1) Calibrated instrumentation should be installed to record the maximum loads or maximum vertical load factor attained during the taxi tests. (2) The taxi surface should be evaluated for compliance with the rule. Corrugated surfaces as well as broken or uneven surfaces (in accordance with the rule) should be used. (3) Representative typical taxi speeds, up to the maximum selected by the applicant, should be attained over the selected taxi surfaces. (4) A light and heavy rotorcraft we ight condition should be evaluated. (5) Limitations appropriate for the ro torcraft design should be included in the flight manual. If these tests indicate that it is unlikely that limit load factors will be attained while taxiing, flight manual limitations may not be necessary. (6) Pertinent taxi information obtai ned from these test conditions may be included in normal procedures of the flight manual. AC 27.239. § 27.239 SPRAY CHARACTERISTICS . a. Explanation . The intent of this requirement is to evaluate by demonstration that water spray does not obscure visibility (day or night) or damage the rotorcraft during normal waterborne operation (for those rotorcraft which have waterborne or amphibious capability). b. Procedures . (1) The following maneuvers should be evaluated in ambient conditions up to the proposed sea state or wave height for operation. Page B - 62 ### Page 100 Config. Condition Weight CG Rotor RPM Altitude Remarks 1 Taxi Max Optional Max SL Speeds up to maximum proposed for water operation. 2 Hover Max Opt Max - Determine critical hover height, if any. 3 Takeoff Max Opt Max SL Unstick at maximum proposed water operation speed. 4 Land Max Opt Max SL Touchdown at maximum proposed for water operation. 5 Shutdown Opt Opt - SL Shut down the rotorcraft. 6 Start Max Opt Max SL Start engines and release rotor brake. (2) The maximum sea state or wave height evaluated under this rule should be stated and included in the limitations section of the flight manual. (3) The effect of saltwater contam ination and deterioration of turbine engines and other component parts of the rotorcraft should be considered in accordance with § 27.609 and paragraph AC 27.609. Information on saltwater effect and attendant corrective action should be provided in the flight manual, if appropriate, and in the maintenance manual. AC 27.241. § 27.241 GROUND RESONANCE . a. Explanation . (1) The rule states: “The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning.” This rule is a flight requirement that pertains to demonstrating freedom from dangerous oscillations on the ground. CAR Part 6, predecessor to FAR Part 27, originally contained a “strength requirement” under § 6.203 requiring ground vibration tests. These tests would identify critical vibration frequencies and modes of the rotorcraft. CAR Part 6, Amendment 6-4, effective October 1, 1959, removed this ground vibration requirement because the agency concluded that if any major component has a natural frequency which could be excited by some operating parameter, such a condition would be revealed in the course of other ground and flight tests. The FAA/AUTHORITY apparently was depending on demonstrations under § 6.131/§ 27.241 and the flight load survey data (§ 27.571) to satisfy the objective of the vibration test. However, Part 27, Amendment 27-2, Page B - 63 ### Page 101 contained new § 27.663 adding reliability and damping action investigation requirements for ground resonance prevention means. A ground vibration survey was not reinstituted by the adoption of § 27.663. Compliance with § 27.663 does require investigation and substantiation as stated. (2) “Ground resonance” is a mechanica l instability of the aircraft while in contact with the ground, often when partially airborne. Stated another way “ground resonance” is a self-excited mechanical instability that involves coupling between the in-plane motion of the rotor blade and the motion of the rotorcraft as a whole on its landing gear (reference “Aerodynamic of the Helicopter,” Gessow & Myers, page 308). It is caused by the motion of the blade in the plane of rotation (called in-plane vibration) coupled with a rocking or vertical motion of the aircraft as a whole. The tires, landing gear, and rotor pylon restraint structure act as a spring with a vibration frequency which coincides or couples with the natural in-plane frequency of the blade about a real or effective drag hinge in the plane of rotation. When the frequencies of the two motions (rotor and airframe) approach each other and couple, a violent shaking of the rotorcraft may occur which, if undamped, could result in the destruction of the rotorcraft. (3) Ground resonance can occur due to fl exibility in the rotor pylon restraint system as well as with landing gear flexibilities. This mode of vibration or resonance can happen in flight (called air resonance) as well as on the ground and should be addressed in the certification program. The evaluation should include variations in stiffness and damping that could occur in service to the rotor pylon restraints. (4) Ground resonance may be prevented by placing the first order in-plane vibration frequency above the rotor turning speed. (5) For such configurations which ar e not susceptible to ground resonance (first order in-plane frequency above rotor turning speed), a simple rotor RPM run-up and run-down with appropriate cyclic control displacement (i.e., excitation of any inherent vibrations) is adequate demonstration that a ground resonance condition does not exist. Unhinged “rigid” rotors, such as Bell Helicopter two-blade designs, are this type of rotor system. (6) For configurations that are su sceptible to ground resonance (i.e., first in-plane frequency is below the rotor turning speed), ground resonance is generally prevented by dampers on the blade acting in the plane of rotation, dampers on the landing gear (sometimes serving as oleo struts), or proper placement of the landing gear frequencies combined with rotor and/or landing gear dampers. (7) Elastomeric components (in the rotor pylon support system, possibly in the landing gear, and possibly in the rotor head) are significantly affected by ambient temperature prior to warm-up. Their damping characteristics require thorough investigation for the range of rotorcraft operating environment as noted in § 27.663. b. Procedures . Page B - 64 ### Page 102 (1) Under all conditions, any oscillations which may be introduced should be damped. However, no instability should occur at any operating condition such as during RPM changes from minimum to maximum and idle to maximum. For rotorcraft with wheel gear, uneven taxi surfaces in conjunction with particular taxi speeds, may excite ground resonance and should be evaluated by taxiing on typical surfaces. This evaluation may be conducted in conjunction with the tests of § 27.235. In operation, the resonance characteristics should be checked during takeoff and landing at zero speed and during run-on landings using various power values. (2) For those aircraft equipped with St ability Augmentation Systems (SAS), all ground resonance investigations should be conducted with SAS on and SAS off. This includes the hovering and running takeoffs and landings, taxi tests, and specific ground resonance tests noted herein. Consideration should be given to conducting tests in various SAS configurations such as roll channel on and pitch channel off, where such configurations are possible and authorized. (3) For each rotorcraft configuration te sted, the aircraft should be positioned on the ground in flat pitch with the rotor stabilized at the minimum practical rotational speed or optionally at a speed shown analytically to have significant margin from indicated resonant conditions. Control system inputs should be used to disturb the system for evaluation of subsequent damping. (4) For each incremental increase in rotor speed and for each rotor speed setting at increments of collective pitch settings, cyclic and collective inputs should be investigated prior to proceeding to the next rotor speed setting. These inputs should cover the appropriate range and combinations of amplitude and frequency. The collective pitch setting increments should range from flat pitch to light on the landing gear prior to fully airborne, depending upon the test sequence for minimum risk. (5) Cyclic pitch inputs should be made ei ther by the pilot through the cyclic stick or through a signal-generating device working in conjunction with the cyclic controls. For each frequency of input, amplitude of the inputs should be increased incrementally and ultimately should be large enough to generate responses representative of normal ground and flight operation on the rotor and support system. The inputs should continue for a time sufficient to obtain representative responses, typically time sufficient to execute five complete circles of the cyclic stick (about neutral) at the selected frequency. (6) The excitation frequency should be such as to excite the blade in-plane frequency. Rotor speed settings should be increased to 1.05 times the maximum power-on rotor speed. Collective pitch settings should be increased in increments of not more than 20 percent to maximum collective or alternately to the collective setting required to become partially airborne (when the cyclic is displaced as noted). Page B - 65 ### Page 103 (7) Typically, articulated rotor aircra ft have natural frequencies on the blade in lag of approximately 0.3 times the power-on main rotor RPM. Soft in-plane rotors have natural frequencies approximately 0.7 times the main rotor RPM. Therefore, for example, for a rotorcraft with an in-plane frequency of 0.3/rev, operating at 300 RPM, and with 6 inches of total lateral cyclic stick displacement, the stick should be rotated for 5 revolutions in a 0.6-inch-diameter circle at ((1-.03) x 300 RPM) or 3.5 cycles per second to attempt excitation of possible resonant frequencies. At the conclusion of the excitation, the cyclic stick should be returned to the neutral position while continuing the recording of data listed in paragraph b(13). (8) The excitation process should in clude cyclic excitation inputs from the directional and longitudinal controls if critical for the type of rotorcraft being evaluated. (9) If onset of ground resonance is encountered, one possible corrective action is to increase the collective pitch and rotor speed and become airborne. However, lowering the collective pitch and applying the rotor brake (if installed) or rolling off the throttles has been effective for some designs and is considered a satisfactory procedure if resonance can be consistently stopped. (10) With the rotor speed stabilized, landing should be made at a touchdown speed which minimizes risk. (11) Special Considerations . (i) The influence of variables, including environmental effects, corresponding aircraft component characteristic changes, operational parameters, and surface conditions should be investigated over the ranges proposed for certification. Additionally, the potential of misservicing and possible failure modes should be evaluated. For ground resonance qualification, where practical, variations from the baseline test configuration may be accomplished by ground run (§ 27.663(b) requires investigation of probable ranges of damping), analyses, component tests, aircraft shake test, the specification of special operational procedures in the rotorcraft flight manual, or a combination thereof. Detailed and rational analyses showing acceptable correlation to the baseline tests, and for which the input parameters were verified by drawings, calculations, component static or dynamic tests, or by aircraft shake tests simulating the conditions/configurations in question, may be used to limit testing to only those variables and operational conditions showing marginal or unacceptable system damping. All operational limitations should be clearly stated in the rotorcraft flight manual. A report of the analytical results and/or test results should be submitted per § 27.663. (ii) Potential instability while ai rborne, called “air resonance,” may occur due to the dynamic coupling of the rotor flexibility and the pylon restraint flexibility. The same considerations apply to air resonance as to ground resonance except that the pylon restraint variables replace the landing gear variables. Air resonance should be addressed in the certification program. Page B - 66 ### Page 104 (iii) When operating on the ground, there may be a tendency for the aircraft to exhibit a “ground bounce.” For many configurations, this is a benign, although undesirable phenomenon which may be aggravated by pilot induced oscillations (PIO), particularly if there is little or no friction on the collective. (12) Rotorcraft with fully articulated rotor heads and landing gear oleos in either skid or wheel configuration have tendencies for ground bounce to occur when light on the oleos, either just prior to takeoff, just after landing contact, or during a power assurance check. This bounce may induce ground resonance, particularly if the intensity of the bounce is aggravated by PIO. The corrective action is either to lift off to a hover or to positively lower the collective and remain on the ground.. (13) Instrumentation and Data Acquisition . (i) Atmospheric Conditions (to be manually noted) : (A) Altitude. (B) OAT. (C) Wind velocity. (ii) Aircraft Configuration (to be manually noted) : (A) Gross weight. (B) C.G. (C) Tire pressure. (D) Landing gear oleo pressure. (iii) Instrumentation (for recording during test). (A) Main rotor RPM. (B) Time history of cyclic contro l fore-and-aft and lateral stick position. (C) Time history of collective control stick position. (D) Time history of rotor damper motion.* (E) Time history of pylon component motion.* (F) Time history of landing gear (oleo) motion.* (G) Time history of aircraft motions.* *As required to obtain modal damping Page B - 67 ### Page 105 **SUBPART B - FLIGHT** **MISCELLANEOUS FLIGHT REQUIREMENTS** AC 27.251. § 27.251 VIBRATION . a. Explanation . (1) Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition (rule statement). (2) This flight requirement may be both a qualitative and quantitative flight evaluation. Section 27.571(a) contains the flight load survey requirement that results in accumulation of vibration quantitative data. Section 27.629 generally requires quantitative data to show freedom from flutter for each part of the rotorcraft including control or stabilizing surfaces and rotors. (3) Review Case No. 70 (referenc e FAA Order 8110.6) contains a policy statement concerning compliance with this rule. This policy statement is condensed here for convenience: “The rotorcraft must be capable of attaining a 30° bank angle (turn), at V NE, with maximum continuous power (maximum continuous torque) without encountering excessive roughness/vibration. The FAA/AUTHORITY requires the maneuver demonstration to provide the pilot with some maneuver capability at VNE and further to provide the pilot some margin away from roughness when operating in turbulence.” (This maneuver may result in a descent or a climb.) (4) Section 27.1505 pertains to V NE determination. Section 27.1509 pertains to rotor speed limits determination. b. Procedures . (1) During the company flight test program, the rotorcraft is flown to the appropriate rotor and airspeed limits at several weights to prove that the rotorcraft is free from excessive vibration under appropriate speed, power, and weight conditions. The flight loads survey quantitative data (reference § 27.571) and the applicant’s qualitative and quantitative flight test data must also prove compliance with the requirement prior to issuing an authorization for official FAA/AUTHORITY flight tests. (2) The flight load survey dat a obtained under § 27.571(a) will contain measured data concerning proof of freedom from flutter and excessive vibration. Pertinent critical flight conditions will be reinvestigated during FAA/AUTHORITY flight tests. The specific condition or conditions necessary to demonstrate compliance with § 27.251 vary with the rotorcraft design and with the minimum and maximum rotor Page B - 68 ### Page 106 speeds, VNE and VD speeds, and weight and CG position. An illustration of the speed and RPM demonstration is shown in figure AC 27.251-1. (Also see paragraph AC 27.251b(4).) (3) The airspeed and rotor speed limits investigated and established under §§ 27.33, 27.1503, 27.1505, and 27.1509 are also investigated and made a matter of record in the flight loads survey data. During the official FAA/AUTHORITY/TIA flight tests, critical parts of the rotorcraft may have limited instrumentation to reinvestigate and confirm that the critical conditions investigated during the flight load survey are satisfactory and do not result in excessive vibration. Use of instrumentation is optional if the flight loads data are conclusive. (4) FAA/AUTHORITY policy for certific ation (Review Case No. 70) requires a “rotor roughness” flight demonstration of a 30° bank angle left and right at maximum continuous power (MCP) (maximum continuous torque which may be in excess of the maximum continuous temperature limit) at VNE. To provide the pilot with some margin from roughness, the FAA/AUTHORITY requires maneuver demonstrations of 30° banked turns at VNE without encountering excessive roughness. The maneuver should be conducted with the rotor speed at the minimum RPM and maximum RPM limits. During the flight load survey, this condition should be investigated and data recorded to ensure hazardous loads are not encountered for this “unusual” condition. As indicated, the flight condition will be reinvestigated during the FAA/AUTHORITY flight tests. See paragraph AC 27.251b(2) for illustration of this speed and RPM demonstration. Page B - 69 ### Page 107 (OCR) 0° Bank, Autorotation Level Forward Flight, 105% Max RPM Power-Off g 4 Power-On 30° Bank, Power-On Max RPM Power-Off 02 Max RPM Power-On o? 4 6 ROTOR RPM —Min RPM Power-On ш 8 45 Min RPM Power-Off 01 95% Min REM Power-Off 03 VNE 1.11 VNE **AIRSPEED** 1. Autorotation at 1.11 VNE (AR) at minimum placard rotor speed. 2. Autorotation at 1.11 NNE (AR) at maximum placard rotor speed. 3. Autorotation at NNE (AR) at power-off minimum design limit rotor speed 4. Autorotation at NNE (AR) at power-off maximum design limit rotor speed 5. Forward flight 1.11 VE at minimum power-on rotor speed. 6. Forward flight 1.11 VNE at maximum power-on rotor speed. 7. Right and left turn at VE at maximum power-on rotor speed with 30° bank angle. 8. Right and left turn at VE at minimum power-on rotor speed with 30° bank angle. Note: VNE (AR) may be less than VNE FIGURE 27.251-1 DEMONSTRATION POINTS Page B - 70 ### Page 108 Page B - 71 ### Page 109 Page C - 1 CHAPTER 2. PART 27 **AIRWORTHINESS STANDARDS** **NORMAL CATEGORY ROTORCRAFT** **SUBPART C - STRENGTH REQUIREMENTS** **STRENGTH REQUIREMENTS - GENERAL** AC 27.301. § 27.301 LOADS . a. Explanation . (1) The rule is a general statement concerning limit and ultimate loads and the application of these loads to the rotorcraft. (2) Ultimate loads are limit loads multiplied by the prescribed factors of safety. (3) The specified loads are specif ied to be distributed appropriately or conservatively and significant changes in distribution of the loads, as a result of deflection, are specified to be taken into account. b. Procedures . The design criteria report and/or design loads report should contain data that comply with the rule. AC 27.303. § 27.303 FACTOR OF SAFETY . a. Explanation . (1) Unless otherwise provided by Part 27, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will ensure that the design strength of the rotorcraft is greater than the design loads contained in Part 27. (2) Other rules, §§ 27.561(b)(3) and 27. 787(c), specify use of defined ultimate inertial forces for protection of occupants. b. Procedures . (1) The design criteria report and/or design loads report should contain data that include the appropriate factor of safety. (2) The factor of safety multiplies the limit external and inertial loads. The rule does allow the application of this factor to the resulting “limit internal” stresses if it is more conservative. ### Page 110 Page C - 2 AC 27.305. § 27.305 STRENGTH AND DEFORMATION . a. Explanation . (1) This general rule defines, in relati ve terms, allowable deformation for limit and ultimate loads. (2) The structure is required to be able to support, in a static test, ultimate loads for 3 seconds without failure, or dynamic tests simulating actual load application may be used. (3) Section 27.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by structural analysis. See paragraph AC 27.307. b. Procedures . Any test results, static or dynamic, should satisfy the limitations or acceptance criteria contained in the rule. (1) Any test proposals submitted fo r approval that are used to demonstrate compliance with sections of Part 27 should contain the criteria stated in the rule. (2) Any test results reports shall c ontain data and information showing the test results comply with the standard. AC 27.307. § 27.307 (Amendmen t 27-3) PROOF OF STRUCTURE. a. Explanation . (1) The rule requires compliance with the strength and deformation requirements for each critical loading condition. Certain tests must be conducted as specified. Additional tests for new or unusual design features may be required as noted in § 27.307(b)(6). (2) Structural analysis rather than l oad tests may be used only if the structure conforms to those for which experience has shown the structural analysis method to be reliable. b. Procedures. (1) The design criteria and/or design loads report should contain typical or representative loading conditions from which the critical loading conditions will be selected for analytical substantiation in structural (static and fatigue) reports, dynamics (vibration and stability) reports, and in fatigue, static, dynamic, or operational test reports. ### Page 111 Page C - 3 (2) Whenever tests are used or required, a test proposal or plan should be approved prior to the tests. The test article should have received conformity inspections and should have been accepted by the FAA/AUTHORITY for the test. Test fixtures and instrumentation should also be acceptable to the FAA/AUTHORITY (using DERs as appropriate) prior to the start of the test. The quality control office of the applicant or other qualified personnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FAA/AUTHORITY or DER performing this task. The test proposal may be used to define and to authorize the means to accomplish inspection of the test fixtures and instrumentation. Unnecessary drawings such as test fixture details or layering of approvals are not intended or envisaged by this policy. Drawings, sketches, or photographs have been used by the FAA/AUTHORITY to control and to ensure correct location, direction, and magnitude of loads and other critical test parameters. (3) Structural analysis has been accepted for rotorcraft in place of static tests. Generally, the rotorcraft airframe should have natural frequencies remote to predominant rotor excitation sources, including higher harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this vibration. During the flight load measurement program conducted under § 27.571, critical loaded areas or critical joints may be instrumented with strain gages or other stress strain measuring devices. This actual flight data should be compared to the analytical data to verify accuracy. (4) Paragraph (b) of the rule specifies certain tests. Test proposals should be approved prior to conducting official FAA/AUTHORITY tests. Other paragraphs in this advisory circular pertain to those tests. AC 27.307A. § 27.307 (Amendmen t 27-26) PROOF OF STRUCTURE. a. Explanation . Amendment 27-26 adds the requirement to account for the environment to which the structure will be exposed in operation. This change is intended to codify recent FAA/AUTHORITY and industry practices for the consideration of environmental effects in showing “proof of structure.” b. Procedures . All of the policy materials pertaining to this section remains in effect with the following additions: (1) For either tests or an analysis, environmental effects are now explicitly required. Consideration of loss of strength and stiffness of metals with elevated temperatures and loss of strength and stiffness of composite materials from exposure to heat, moisture, or other operational environments is now required and should be documented in analyses and test reports. (2) MIL-HDBK-5F; AC 20-107B; or MIL- HDBK-17, Vol. I, Rev 1E; Vol. II, Rev. D; Vol. III, Rev E (or later versions) are acceptable sources of data and procedures ### Page 112 Page C - 4 to show compliance with environmental effects of metallic and composite materials, respectively. AC 27.309. § 27.309 DESIGN LIMITATIONS . a. Explanation . (1) The rule requires an orderly se lection and presentation of the basic structural design limitations of the rotorcraft. The applicant is required to establish these structural limitations to facilitate design of the rotorcraft. (2) Refer to the rule fo r the specific requirements. b. Procedures . (1) The design criteria and/or desig n load report should contain the design limits specified. (2) These items are structural design lim its. Other requirements may result in narrowing the ranges of type design limits or in reducing limits. It is not necessary to revise structural design criteria limits to agree with more conservative operational limits established during the certification program. The operational limits may be subsequently expanded by additional flight tests to agree with design limits. ### Page 113 Page C - 5 **SUBPART C - STRENGTH REQUIREMENTS** **FLIGHT LOADS** AC 27.321 § 27.321 (Amendmen t 27-11) FLIGHT LOADS--GENERAL. a. Explanation . (1) The rule specifies the way the lo ads will be applied to the rotorcraft. It requires load analysis from minimum to maximum design weight. Any practical distribution of disposable loads must be included in the analysis. (2) Paragraph (a) of the rule states: “The flight load factor must be assumed to act normal to the longitudinal axis of the rotorcraft, and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the center of gravity.” b. Procedures . (1) Derivation of the flight loads is required by and specified in §§ 27.337 through 27.351. This rule requires flight load determination from minimum to maximum weight and for disposable loads. (2) The application of the design loads derived from the flight load factor will be as specified. The flight loads analysis data must comply with the rule. AC 27.337 § 27.337 (Amendment 27-26) LIMIT MANEUVERING LOAD FACTOR. a. Explanation . The rotorcraft must be designed and substantiated to load factors as specified to provide a minimum level of structural integrity of the rotorcraft airframe and rotors. (1) A range of design positive load fact ors from +3.5 to +2.0 may be used. (2) A range of design negative load fact ors from -1.0 to -0.5 may be used. (3) Load factors inside the range of +3 .5 to -1.0 may be used provided the probability of exceeding the design load factors is shown by analysis and flight tests to be extremely remote and the selected load factors are appropriate to each weight condition between design maximum and minimum weight. (4) Load factors exceeding these “minimums” may be used. b. Procedures . ### Page 114 Page C - 6 (1) The applicant may elect to subs tantiate the rotorcraft for a design maneuvering load factor less than +3.5 and more than -1.0. Whenever this option is used, an analytical study and flight demonstration are required. (i) The maximum positive design load factor of +3.5 is generally at a weight below maximum gross weight. The maximum thrust capability of the main rotor, combined with incremental lift of wings or sponsons, if installed, results in a maximum design positive load factor. An example of a load factor-gross weight curve is shown in figure AC 27.337-1. Note the minimum positive design load factor is +2.0 even though the required analysis and flight demonstration may prove the rotorcraft is not capable of achieving this load factor. This curve also illustrates compliance with § 27.321(b)(1) since the design load factor varies with gross weight. (ii) The largest negative design lo ad factor is -1.0; however, several current rotorcraft designs are not capable of achieving a negative load factor. Therefore, -0.5 has been an acceptable structural design negative load factor for certain rotorcraft designs. (2) Whenever the applicant analytically substantiates the lower load factors allowed by § 27.337(b), the flight demonstration required by § 27.337(b) must be conducted. The flight test personnel should determine that the demonstration shows the probability of exceeding the selected design load factors (those factors less than +3.5 and more than -1.0) is extremely remote. (See Order 8110.4, paragraph 166c(2)(c)). (3) A numerical value has not been assigned to “extremely remote” in this standard. ### Page 115 (OCR) +3.5 _ Represents curve of Design thrust + lift Load Factor Nz - +2.0 - 1.0 - Minimum Gross Weight Lbs. Maximum LOAD FACTOR GROSS WEIGHT CURVE FIGURE 27.337-1 Page C - 7 ### Page 116 Page C - 8 AC 27.339 § 27.339 (Amendment 27- 11) RESULTANT LIMIT MANEUVERING L O A D S . a. Explanation . The rule specifies or defines the application of rotor and lift surface loads to the rotorcraft. (1) The design maneuvering load factors required by § 27.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift. (2) Sections 27.321, 27.337, and 27.341 all complement one another and result in the derivation of design flight loads that will be imposed to ensure structural integrity of the rotorcraft. (3) The following assumptions and conditions are specified in the rule. (i) The rule requires application of appropriate loads at each rotor hub and auxiliary lifting surface. (ii) Power-on and power-off flight with maximum design rotor tip speed ratio and specific conditions that must be considered. (iii) Rotor tip speed ratio, defi ned in the rule, has been carried forward from the initial rotorcraft certification rules issued in 1946. The rotor tip speed ratio is a basic parameter used in calculating rotor aerodynamic forces. b. Procedures . (1) The rule specifies an acceptable a ssumption concerning application of the rotorcraft maneuvering loads. (2) The rotor tip speed ratio is a par ameter found in textbooks and other books such as NACA Report No. 716. The equation in the rule contains angle “a.” Report No. 716 also defines angle “a” as the angle of attack of the rotor disk. This definition is more easily understood than the definition contained in the rule. (3) The rotorcraft design loads are derived as prescribed by §§ 27.321, 27.337, and 27.341. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule. AC 27.341 § 27.341 GUST LOADS . a. Explanation . (1) The rotorcraft must be substantia ted for the loads derived from 30 feet per second vertical gusts from hovering to 1.11 VNE (i.e., VD). ### Page 117 Page C - 9 (2) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward. b. Procedures . (1) Either sharp-edged (instantaneous) gusts or sharp-edged gusts modified by an alleviation (attenuation) factor may be used for calculating aerodynamic loads for the rotorcraft and any installed stabilizing surfaces. The following conditions may be used: (i) Vertical gusts may be consider ed normal to the flight path of the rotorcraft except during hover or low speed flight (20 knots or less) when the gusts may be assumed normal to the longitudinal axis of the rotorcraft. (ii) A primary effect of encounter ing the gust is to change the lift of the rotors and rotorcraft surfaces. Of primary concern is the gust load or lift created by the main rotor or rotors. The lift increment of the horizontal stabilizing surface and fuselage is generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis. (iii) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust. (iv) The gust velocity may be assumed uniform across the rotorcraft. (v) Gust loads on the stabilizing surfaces are required as stated in paragraph AC 27.413. (2) The rotorcraft design maneuvering load factors may generally exceed the design gust load factors calculated in compliance with this rule. This may be attributed to the small incremental change in lift due to the 30 FPS gust. Nonetheless, design gust loads for the rotorcraft shall be calculated as specified in the rule to ensure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load factor. (3) For further information about rotorc raft gust response characteristics, see Paper No. 9 presented at the AHS/NASA -Ames Specialist’s Meeting on Rotorcraft Dynamics, February 13-15, 1974. The paper, entitled, “Helicopter Gust Response Characteristics Including Unsteady Aerodynamics Stall Effects,” was written by P.J. Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also. AC 27.351. § 27.351 (Amendment 27-26) YAWING CONDITIONS. a. Definitions. ### Page 118 AC 27-1B, Chg 2 4/25/06 Page C - 10 (1) Suddenly . For the purpose of this section, ‘suddenly’ is defined as an interval not to exceed 0.2 seconds for complete control input. A rational analysis may be used to substantiate an alternative value. (2) Zero Yaw . Normal, 1-g, level flight condition with either zero bank angle or zero sideslip. b. b. Explanation . The rule requires a rotorcraft “structural” yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed, or hover, to VH or VNE, whichever is less. The rotorcraft must be structurally safe for the thrust capability of the directional control system. (1) The rotorcraft structure must be designed to withstand the loads for the specified yawing conditions. The standard does not require a structural flight demonstration. It is a structural design standard. (2) This standard applies only to pow er-on conditions. Autorotations need not be considered. (3) This standard requires the maximum allowable rotor RPM consistent with the flight conditions, including special operational rotor settings. (4) For the purposes of this sect ion, the analysis may be performed at international standard atmosphere (ISA) sea level conditions. (5) The rotorcraft structure must be designed to withstand the loads for the specified sideslip conditions. This includes, but is not limited to: (i) Main cabin, tailboom, and vertical control surfaces . ( i i) Tail rotor structures, including t he fitting attachments to the frame. (iii) Windows, doors, and other transparencies. (iv) Landing gear and retracting mechanism. (v) Fairings and cowlings. (6) Maximum displacement of the direct ional control, except as limited by pilot effort (§ 27.397(a)), is required for the conditions cited in the rule. Control system limiting devices may be used, however the probability of failure or malfunction of these system(s) should be considered (see Figure AC 27.351-2). This evaluation may include Flight Manual Limitations, if failure of the system is reliably indicated to the crew. (7) Both right and left yaw conditions should be evaluated. ### Page 119 Page C - 11 (8 ) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust (superpositioning). (9) Loads associated with sideslip angl es exceeding the values of Figure AC 27.351-1 do not need to be considered. The corresponding points of the maneuver may be deleted. c. Procedure . The design loads should be evaluated within the limits of Figure AC 27.351-1 or the maximum capability of the rotorcraft, whichever is less; at speeds from zero to VH or VNE, whichever is less, for the following phases of the maneuver: (1) With the rotorcraft at an initial tr im condition (1 g level flight and zero yaw), the cockpit directional control is suddenly displaced to the maximum deflection limited by the control stops or by the maximum pilot force specified in § 27.397(a). This is intended to generate a high tail rotor thrust. (2) While maintaining maximum cockpit directional control deflection, within the limitation specified in c(1) of this AC paragraph, allow the rotorcraft to yaw to the maximum transient sideslip angle or to the value defined in Figure AC 27.351-1, whichever is less. This is intended to generate high aerodynamic loads. (3) Allow the rotorcraft to stabilize at the maximum steady-state sideslip angle. In the event that the maximum steady state angle is greater than the value defined in Figure AC 27.351-1, the rotorcraft should be trimmed to the value of the angle using less than maximum cockpit directional control deflection. (4) With the rotorcraft yawed to the st atic equilibrium sideslip angle specified in c(3) of this AC paragraph, the cockpit control is suddenly returned to its initial trim position. This is intended to combine a high tail rotor thrust and high aerodynamic restoring forces. ### Page 120 AC 27-1B, Chg 2 4/25/06 Page C - 12 Figure AC 27.351-1 Safety Factors for Probability of Failure 15 90 0 0.6 Vne Vh or Vne whichever is less **AIRSPEED** SIDESLIP ANGLE ( deg. ) ### Page 121 Page C - 13 • For static strength substantiation, each part of the structure should be able to withstand without failure, the loads generated by the maneuver described in the rule multiplied by a factor of safety depending on the probability of being in this failure state. The factor of safety is defined in the figure below: Qj = (T j)(Pj) where: T j = Average flight time spent with a failed control limiting system j (in hours) P j = Probability of occurrence of failure of the control limiting system j (per hour) Note: If Pj is greater than 10-3 per flight hour then a 1.5 factor of safety should be applied to all limit load conditions specified in this standard. Figure AC 27.351-2 ### Page 122 Page C - 14 AC 27.361 § 27.361 (Am endment 27-23) ENGINE TORQUE. a. Explanation . (1) The rotorcraft shall be designed fo r limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. Amendment 27-23 separated the standard into paragraphs for turbine and reciprocating engine limit torque values. (2) Turbine engine limit torque fo r design purposes (Amendment 27-23) was redefined into four cases and the torque values determined will be used. For example, sudden engine stoppage is introduced as one of the cases which is applied to the engine and the engine suspension and restraint system. Emergency operation of governor-controlled turboshaft engines is another case. (3) Torque factors are also specifi ed for reciprocating engines having two or more cylinders in paragraph (b) of the standard. (4) Sections 27.547(e)(1)(ii) and 27.549(d) , respectively, refer to the application of engine torque to design of main rotor structure and engine mount and adjacent structure. b. Procedures . (1) The engine torque associated wit h the maximum continuous (MC) power condition for reciprocating engines should be multiplied by the appropriate torque factor to obtain the limit engine torque value used for structural substantiation of the rotorcraft. (2) The torque values associated wit h MC power at the minimum power-on RPM limit should be used. Maximum power-on speed limit will result in a lower torque value when calculating torque from design horsepower values. However, due to piston engine power output characteristics, an engine may produce a higher torque at higher engine speeds contrary to the previous statement. The torque factor should account for this characteristic. (3) Turbine engine limit torque valu es are determined for the four cases specified. Two cases are related to the endurance test of §§ 27.923 and 27.927. (4) For sudden stoppage of turbine engines the engine manufacturers can reasonably provide FAA/AUTHORITY approved data to the applicant on inertia of rotating parts and the deceleration time expected in the event of sudden engine stoppage. This condition usually generates critical loads in the engine mounting and restraint system. These manufacturer’s data should be acceptable for use in compliance with this part of the standard. ### Page 123 Page C - 15 **SUBPART C - STRENGTH REQUIREMENTS** **CONTROL SURFACE AND SYSTEM LOADS** AC 27.391. § 27.391 CONTROL SURFACE AND SYSTEM LOADS--GENERAL. a. Explanation . This general standard concerns requirements for design loads of tail rotors, control or stabilizing surfaces, and their control system. b. Procedures . The design criteria and/or the design loads report shall contain the loads dictated by the referenced rules. (See paragraphs AC 27.395, AC 27.397, AC 27.399, AC 27.401, AC 27.403, AC 27.411, and AC 27.413.) AC 27.391A. § 27.391 (A mendment 27-26) GENERAL. a. Explanation . Amendment 27-26 adds an explicit reference to § 27.427, Unsymmetrical Loads (paragraph AC 27.427), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 27.399, Dual Control System (paragraph AC 27.399), is also added for clarification. In addition, §§ 27.401, 27.403, and 27.413 were removed by this amendment since these references and requirements were adequately addressed in other standards. b. Procedures . The referenced AC paragraphs become paragraphs AC 27.395, AC 27.397, AC 27.399, AC 27.411, and AC 27.427. AC 27.395. § 27.395 CONTROL SYSTEM . a. Explanation . Control system design loads and the application of these loads are contained in this rule. (1) Paragraph (a) of the rule specif ies the way or means of reacting the minimum design loads specified in §§ 27.397 and 27.399 (for dual control systems). Except reduced design loads, not less than 0.60 of those specified in §§ 27.397 and 27.399 for dual control system, may be used as specified. The standard also applies to those control systems that may have more than one stop in a system. The design loads must be imposed on the system from the pilot’s control to any stop in the control system. (2) Minimum design loads imposed on the control system from a stop to a rotor blade or a control surface or device shall be: (i) The maximum pilot forces obtainable in normal operation; and ### Page 124 Page C - 16 (ii) If low operational loads may be exceeded as noted in § 27.395(b)(2), the system shall support without yielding 0.60 of the loads specified in §§ 27.397 and 27.399 for dual control systems. (3) Section 27.695 concerns standards for a power boost and power-operated control system. This standard, in effect, imposes a fail-safe standard for hydraulic aspects of a control system. Where appropriate to a particular design, the control system must therefore sustain without yielding, the maximum output force of the actuator when complying with § 27.395(a). The pilot input forces are not added to the actuator output forces according to this standard for normal category rotorcraft. These forces are independently applied to the control system. (4) Control system design features and tests requirements are found in §§ 27.619 and 27.625, respectively. Special factors such as casting, bearing, and fitting factors that may be appropriate for the design are contained in §§ 27.619 and 27.625, respectively. b. Procedures . (1) The design criteria and/or a desi gn loads report that includes the primary control system design loads should be submitted for FAA/AUTHORITY approval. (2) The rotorcraft control system may be tested to ultimate design loads or may be analyzed for the ultimate design loads. See paragraph AC 27.307. (i) A static test proposal for testing the control system to show compliance with the rules should be approved before conducting the test. Where compliance is to be determined by tests, limit load tests, as discussed in paragraph AC 27.681, and/or ultimate load tests may be performed. Test results shall be documented. (ii) If tests are not c onducted, a structural analysis of the control system is required. Appropriate factors from §§ 27.623 and 27.625 must be used as specified. Tests may not be required when adequate similarity of systems and support structure is determined and where adequate structural analysis is furnished. (3) If a part of the control system is no t stiff or rigid enough to react the design loads specified in §§ 27.397 and 27.399, that part of the system may be substantiated for lower loads as prescribed. (i) The limit design loads are those loads specified in §§ 27.397 and 27.399; (ii) The maximum that can be obt ained in normal operation and that is allowed by the system; except ### Page 125 Page C - 17 (iii) The limit design loads may not be less than 0.60 of the limit pilot forces specified. (iv) For example, if a small control surface or servo tab is lightly loaded, its control system must be stiff enough to react the control surface loads and to provide surface deflection to control the rotorcraft. The normal operational loads may be very low, such as 10 pounds maximum. Nonetheless, the design limit load shall be 0.60 times the limit single pilot forces specified in § 27.397. Note that the system must not yield under these loads. (v) For example, if a dual but pr imary manual control system such as a tail rotor control is lightly loaded, the control system, from the stops to the rotor blades, may be designed for minimum loads equal to 0.60 times the limit dual pilot forces specified in § 27.399. (vi) If a power actuator is a part of a rotor control system, the design limit force for the affected parts shall be the maximum output force of the actuator at any operational condition (including any load/pressure after a single failure in the hydraulic system). (4) Controls proof and operation test is required by §§ 27.307(b)(2) and (b)(3), 27.681, and 27.683. This test is conducted using the design limit loads approved under § 27.395. (See paragraphs AC 27.681 and AC 27.683.) AC 27.395A. § 27.395 (Am endment 27-26) CONTROL SYSTEM. a. Explanation . Amendment 27-26 extensively rewrites § 27.395(b) to more clearly incorporate design condition loads for typical powered control systems. New requirements include substantiation for loads resulting from “each normally energized power device, including any single power boost or power activator failure in the control system.” There are also new minimum loads for control system designs in which operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. The old loads were 0.60 times the limit pilot forces of § 27.397; the new loads are 100 percent of the limit pilot forces specified in § 27.397. b. Procedures . The procedures of paragraph AC 27.395 continue to apply except that the increased loads in § 27.395(b)(4) of 100 percent of limit pilot forces are specified for systems where operational loads may be exceeded by jamming, ground gusts, control inertia, or friction. AC 27.397. § 27.397 (Amendment 27-11) LIMIT PILOT FORCES AND TORQUES . a. Explanation . Design forces are contained in the rule. ### Page 126 Page C - 18 (1) Primary controls, pilot and copilo t, should be designed for the limit pilot forces specified in paragraph (a) of the rule unless higher forces are used. (2) For other operating controls, such as flap, tab, stabilizer, rotor brake, and landing gear, design limit forces are specified in paragraph (b). b. Procedures . (1) Design loads specified in the rule may be used in required structural tests and in any structural strength analysis of the control systems submitted in compliance with other rules. (2) Operation tests of the control systems noted in other rules require application of these forces also. AC 27.399. § 27.399 DUAL CONTROL SYSTEM . a. Explanation . Design limit loads are specified for dual control systems. Pilot effort forces applied in opposition and in the same direction are required for dual control systems. b. Procedures . (1) Design loads specified in the rule may be used in required structural tests and in any structural strength analysis submitted for compliance with the other rules. (2) Operation tests of the control systems, noted in other rules, require application of these forces also. AC 27.401. § 27.401 (Amendment 27- 3) AUXILIARY ROTOR ASSEMBLIES. a. Explanation . (1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by § 27.923, and structural strength substantiation is required. Section 27.401(b) specifically refers to structural strength substantiation of detachable blade systems for centrifugal loads resulting from maximum design rotor RPM. (2) The rotor blade structure must have sufficient strength to withstand not only aerodynamic loads generated on the blade surface, but also inertial loads arising from centrifugal, coriolis, gyroscopic, and vibratory effects produced by this blade movement. Sufficient stiffness and rigidity must be designed into the blades to prevent excessive deformation and to ensure that the blades will maintain the desired aerodynamic characteristics. As a design objective, the structural strength requirements should be met with the minimum material. Excess blade weight imposes extra centrifugal loads ### Page 127 Page C - 19 that may increase the operating stress levels. Blade weight and strength should be optimized. Even though a structural strength analysis for the blade design loads is required, a flight load survey and fatigue analysis are also required by § 27.571. (3) Section 27.1509 defines the desi gn rotor speed as that providing a 5 percent margin beyond the rotor operating speed limits. b. Procedures . (1) The endurance tests prescribed by §§ 27.923 and 27.927 require achieving certain speeds, power, and control displacement for the auxiliary (tail) rotor as well as the main rotor. The parts must be serviceable at the conclusion of the tests. (2) Structural substantiation of the au xiliary (tail) rotor is required to ensure integrity for the minimum and maximum design rotor speeds and the maximum design rotor thrust in the positive and negative direction. Thrust capability of the rotor should offset the main rotor torque at maximum power as required by § 27.927(b). (i) The maximum and minimum oper ating rotor speed, power-off, is 95 percent of the maximum design speed and 105 percent of the minimum design speed, respectively. (ii) The rotor operating sp eed limits shown during the official FAA/AUTHORITY flight tests must include the noted 5 percent margin with respect to the design speeds. (iii) The auxiliary rotor generally has a positive and negative pitch limit that ensures adequate directional control throughout the operating range of the rotorcraft. The power-off rotor speed limits are generally broader than the power-on rotor speed limits because of the required autorotational rotor speed characteristics. Thus, the auxiliary rotor design conditions concern the maximum and minimum design rotor speeds in conjunction with the maximum positive or negative pitch thrust, as appropriate. Thrust capability and precone angle of the rotor, if any, will significantly influence the rotor design loads. The variations in rotor design features and an example of substantiation would be too lengthy to include here. However, ANC-9, “Aircraft Propeller Handbook” contains principles that may be applied to tail rotor designs. Tail rotors may be considered a special propeller design. (iv) Bearings are generally used in the tail rotor installation to allow flapping and feathering motion of the blades. The bearing manufacturer’s ratings of these bearings must not be exceeded. Bearings generally used in main and tail rotors are classified as ABEC Class 3, 5, or 7. Class 7 is the highest quality presently available. Satisfactory completion of the endurance tests of §§ 27.923 and 27.927 is a means of proving that use of a particular bearing is satisfactory. ### Page 128 Page C - 20 (v) The analysis must include appropri ate special factors, casting factors, bearing factors, and fitting factors prescribed by §§ 27.619, 27.621, 27.623, and 27.625, respectively. The fitting factor of 1.15 must be applied in the analysis of the tail rotor installation. AC 27.401A. § 27.401 (Amendment 27-27) AUXILIARY ROTOR ASSEMBLIES. a. Explanation . Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, and 27.341. b. Procedures . The policy material pertaining to this section is retained as supplemental information. AC 27.403. § 27.403 AUXILIARY ROTOR ATTACHMENT STRUCTURE. a. Explanation . (1) The auxiliary rotor attachment stru cture(s), which is considered to include gearboxes, must be designed to withstand design limit loads that occur in flight and on landing. These design loads that generally consist of the following must be established for the particular flight and landing condition under consideration. (i) Inertia loads generated by li near and angular accelerations of the auxiliary rotors and their gearboxes, combined with-- (ii) Thrust and torque loads developed by the auxiliary rotors. The linear and angular acceleration loads imposed by the weight of the tail rotor and gearbox are generally derived from airframe loads data. Thrust and torque output of the tail rotor are derived during external aerodynamic and landing loads development for pertinent flight and landing conditions. (2) General rules related to proof of structure loads and factor of safety are §§ 27.307, 27.301, 27.303, and 27.305. b. Procedures . (1) The angular and linear accelerati on loads combined with appropriate tail rotor thrust and torque for the critical conditions shall be imposed on the tail rotor gearbox mount lugs, the airframe mounting structure, and the attaching hardware. (2) The yaw and maximum power clim b conditions are generally critical. Landing and maneuvering conditions with and without power may also impose high inertia and rotor thrust and torque loads on the attachment structure. ### Page 129 Page C - 21 (3) The derivation of the loads and conditi ons is too extensive to include here. Additional information can be found in the U.S. Army Material Command Report AMCP 706-201, “Engineering Design Handbook: Helicopter Engineering, Part One, Preliminary Design.” AC 27.403A. § 27.403 (Amendment 27- 27) AUXILIARY ROTOR ATTACHMENT STRUCTURE . a. Explanation . Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, 27.341. b. Procedures . The policy material pertaining to this section is retained as supplemental information. AC 27.411. § 27.411 GROUND CLEA RANCE: TAIL ROTOR GUARD. a. Explanation . (1) The rule requires specific protecti on to prevent the tail rotor from contacting the landing surface during a normal landing if it is possible that the tail rotor will contact the surface. The rule states that it must be impossible for the tail rotor to contact the surface during a normal landing. (2) If a guard is requir ed, the guard and its supporting structure must withstand suitable design loads. (3) Section 27.501(c)(1) contains skid landing gear drag requirements that may be applied to the guard design loads. b. Procedures . (1) The applicant may submit sketc hes or drawings showing probable clearance with typical level landing surfaces during normal landings. Typical attitudes such as nose-high autorotation, or autorotation with power-on landing, or other possible tail-low attitudes should be investigated. If the drawings or sketches reveal that it is not likely the tail rotor will contact the landing surface, this minimum clearance with the landing surface may be confirmed during official FAA/AUTHORITY flight tests, such as HV and landing tests. The clearance may be confirmed by having a frangible device of suitable length (i.e., a balsa wood dowel) extending beyond the guard and attached to the tail rotor guard or other appropriate fuselage part. If the device is not damaged, broken, or no contact is made with the surface, compliance has been demonstrated. (2) If it is possible for the tail rotor guard to contact the landing surface, suitable design loads must be established for the guard. ANC-2a dated March 1948, “ANC Bulletin Ground Loads,” paragraph 6.4, entitled “Tail Bumper Criteria,” is an acceptable ### Page 130 Page C - 22 means of deriving the rotorcraft kinetic energy that shall be absorbed by the guard. This method is noted here for convenience. (i) The tail rotor guard shall be able to absorb the kinetic energy of the rotorcraft in its most unfavorable CG position in the tail-down landing attitude. The kinetic energy that the tail rotor guard should be capable of absorbing may be determined by the following: W V S2 KY2 KE = X 2g (KY2 + 1B2) where-- V S = vertical speed ft/sec, derived from § 27.725(a) K Y = pitching radius of gyration - ft from pitching axis 1 B = distance from most critical CG location to the guard or bumper contact point - ft W = gross weight less rotor lift from § 27.473(a) - lbs g = 32.2 ft/sec 2 (ii) Other, more recent, analytic al techniques (most utilizing computer programs) may, of course, be used rather than the ANC-2a means after proper substantiation for applicability and validity. (iii) The tail rotor guard should not fail when the limit and ultimate load, which is derived from a combination of the limit kinetic energy and the guard resulting limit deflection required to dissipate the energy, is imposed on the guard and the rotorcraft tail (see § 27.305). (3) Substantiation of the guard, skid, or bumper for the design loads derived may be accomplished by test or analysis as stated in § 27.307(a). (4) Several rotorcraft tail rotor guards are installed solely for the protection of ground personnel from the rotating tail rotor. For guards installed for this purpose, the applicant should use prudent and reasonable design loads and features. Such guards should not present a hazard to the rotorcraft because of its design features. AC 27.413. § 27.413 STABILI ZING AND CONTROL SURFACES. a. Explanation . Minimum design loads are specified for stabilizing as well as control surfaces. (1) Paragraph (a) of the rule require s application of minimum empirical design loads, application of critical maneuvering loads, and application of critical maneuvering loads combined with vertical gust loads (30 feet per second or 17.8 knots per § 27.341). ### Page 131 Page C - 23 (2) Paragraph (b) requires load distribu tions that closely simulate actual pressure distributions. Both spanwise and chordwise distributions are intended. (3) These surfaces are used for st ability and control thereby hopefully extending the CG range and increasing the airspeed of modern designs. (4) To “closely simulate actual pressure condition” on the surfaces, unsymmetrical loads are also required on horizontal surfaces. An arbitrary distribution, if conservative, may be used. (5) It is noted § 27.571 requires fatigue s ubstantiation of the flight structure which will include control and stabilizing surfaces. (6) If the surface is controllable, a proof and operation test of the surface control system is required by §§ 27.681 and 27.683. b. Procedures . Modern rotorcraft designs have generally employed a fixed or a wholly movable, not split or divided, stabilizing or control surface. (1) Design Loads. (i) Limit loads of 15 pounds per square foot will apply up to approximately a 90-knot design airspeed. Above a 90-knot design airspeed (1.11VNE), the coefficient (CN = 0.55) imposes higher limit loads on the surface. The coefficient CN is assumed normal to the chordline of the section. (ii) In addition, combined maneuver ing and vertical up or down gust loads may impose the highest limit loads on the control surfaces of rotorcraft. This is attributed to the change in angle of attack and change in resultant airspeed. (iii) The applicant may choose to derive the limit loads using maximum aerodynamic coefficients for the surface under consideration at the maximum design airspeed combined with a 17.8-knot gust. This would be acceptable provided these design loads exceed the minimum loads derived from a CN = 0.55 at design airspeed or exceed 15 pounds per square foot load on the surface. (2) The load distribution on the surface should closely simulate actual pressure distributions. (i) The spanwise load may be rectangular, or other acceptable conservative distributions may be used. The method developed by O. Schrenk in NACA TM 948, 1940, is an acceptable method for approximation of spanwise distribution. ### Page 132 Page C - 24 NOTE: The method is valid for aspect ra tios of 5 through 12 and for rectangular planforms such as used on rotorcraft, other planforms may be acceptable as prescribed in the TM. (ii) The chordwise distribution appr opriate for the aerodynamic shape or a conservative distribution should be used. (iii) The flight load survey conducted under § 27.571 may be used to confirm design parameters and possible load distribution data. On controllable surfaces, the pitching moment (control loads) may be measured for fatigue substantiation of the control system. The control stabilizing surfaces may be subject to loads measurement and possible fatigue tests for fatigue substantiation also. (3) Proof of the struct ure for the required loads is specified in §§ 27.301, 27.303, 27.305, and 27.307. Tests or analysis may be used as prescribed. If analysis is used, fitting factors and other appropriate factors prescribed by the rules of §§ 27.625, 27.621, and 27.623 will be required in the analysis. AC 27.413A. § 27.413 (Amendment 27-27) STABILIZING AND CONTROL SURFACES . a. Explanation . Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, and 27.341. b. Procedures . The policy material pertaining to this section is retained as supplemental information especially as reference material for paragraph AC 27.341 (§ 27.341). AC 27.427. § 27.427 (Amendment 27-27) UNSYMMETRICAL LOADS. a. Explanation . Amendment 27-26 added the standard and Amendment 27-27 amended it. Minimum unsymmetrical design loads are specified for horizontal tail surfaces and also vertical tail surfaces whenever they support the horizontal tail surfaces. (1) Loads are derived by rational analysis, or for earlier certification bases, the prescribed empirical loads of § 27.413 may be used. Section 27.413 was removed by Amendment 27-27 since the requirements are adequately addressed in §§ 27.337, 27.339, and 27.341. (2) Rational loads, appropriate for the aerodynamic surfaces, should be distributed according to the standard. (3) When vertical tail surfaces support the horizontal tail surfaces, the vertical tail surfaces and supporting surfaces are required to support the critical combination of vertical and horizontal surface loads. ### Page 133 Page C - 25 b. Procedures . Two basic loading conditions are required by § 27.427 for each of the two basic empennage configurations. (1) Horizontal surfaces suppor ted by the tail boom or fuselage. Structural substantiation should be provided for all six combinations shown in figure AC 27.427-1. All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly. (i) First unsymmetric al loading condition: (A) 100 percent of the flight load is applied to one side of the plane of symmetry; and 0 percent of the flight load is applied to the other side of the plane of symmetry. (B) For surfaces with end plates or other similar devices, the load distribution will be changed accordingly. (ii) Second unsymmetr ical loading condition: 50 percent of the flight load is applied to one side of the plane of symmetry acting up; and 50 percent of the flight load is applied to the other side of the plane of symmetry acting down. (2) Horizontal surfaces s upported by a vertical surface. Structural substantiation should be provided for all six combinations shown in figure AC 27.427-2. All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly. (i) First unsymmetric al loading condition: 100 percent of the flight load is applied to one side of the plane of symmetry; and 0 percent of the flight load is applied to the other side of the plane of symmetry. (ii) Second unsymmetr ical loading condition: 50 percent of the flight load is applied to one side of the plane of symmetry acting up; and 50 percent of the flight load is applied to the other side of the plane of symmetry acting down. ### Page 134 (OCR) АС 27-1B 9/30/99 Figure AC 27.427-1 (View Looking Forward) Page C - 26 ### Page 135 (OCR) Figure AC 27.427-2 (View Looking Forward) Page C -27 ### Page 136 Page C - 28 **SUBPART C - STRENGTH REQUIREMENTS** **GROUND LOADS** AC 27.471 § 27.471 GROUND LOADS--GENERAL . a. Explanation . This regulation specifies that limit ground loads must be considered which are: (1) External loads caused by l anding (ground) conditions for skid and wheel landing gear equipped rotorcraft and by ground taxiing loads as specified in § 27.235 for wheel landing gear equipped rotorcraft. (2) Loads considering the rotorc raft structure as a rigid body. (3) Loads in equilibrium wit h linear and angular inertia loads. (4) The critical center of gravity “mu st be selected so that the maximum design loads are obtained in each landing gear element.” b. Procedures . (1) The standards to be considered are specified in §§ 27.473 through 27.505. These associated standards cover landing gear arrangements, landing conditions, and ground loading conditions (for wheel landing gear rotorcraft). (2) Drop tests may be used to veri fy landing load factors. (See paragraph AC 27.725.) (3) The application of the design load s derived from the landing load factors will be as specified for each element affected by landing or ground loading conditions (for wheel landing gear rotorcraft). (4) During the applicant’s flight test program, the landing load factors for skid and wheel landing gear rotorcraft and taxiing load factors for wheel landing gear rotorcraft are monitored to assure the design load factors used are adequate. See paragraph AC 27.235 of this document for § 27.235 policy. AC 27.473 § 27.473 (Amendment 27- 2) GROUND LOADING CONDITIONS AND ASSUMPTIONS . a. Explanation . The rotorcraft is to be designed for the maximum weight. A rotor lift of two-thirds of the design maximum weight may be used. The minimum limit landing load factor is determined by the drop tests of § 27.725. ### Page 137 Page C - 29 b. Procedures . Loads for the landing conditions are derived considering mass (equal to the maximum weight) and rotor lift (equal to two-thirds of the maximum weight) acting through the center of gravity throughout the landing impact. Unbalanced external loads resulting from asymmetric loading conditions are reacted as specified in the individual subparagraphs. The rotorcraft must be substantiated for ultimate landing loads by either test or analysis utilizing an ultimate load factor of 1.5 applied to the limit load factor of not less than that substantiated under § 27.725. AC 27.475 § 27.475 TIR ES AND SHOCK ABSORBERS. a. Explanation . This section specifies the tire and shock absorber position to be used in ground load derivations. b. Procedures . Ground loads are to be derived with the tires in static (1g) position and the shock absorbers “in their most critical position.” The determination of the “most critical position” for the shock absorbers generally requires a load versus deflection test or analysis of the shock absorber system and a determination of the effect of both load and deflections on the shock absorber, attachment structure, and substructure designed by ground loads. AC 27.477 § 27.477 LANDI NG GEAR ARRANGEMENT. a. Explanation . This section specifies the individual standards to be used for ground load conditions for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity. NOTE: Section 27.497 gives ground loading conditions for landing gear with tail wheels, and § 27.501 gives ground loading conditions for landing gear with skids. b. Procedures . The ground loading conditions of §§ 27.235, 27.479 through 27.485, and 27.493 will be used for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity. This includes forward wheels on separate axles. AC 27.479 § 27.479 LEVEL LANDING CONDITIONS . a. Explanation . This section provides explicit level landing load criteria for landing gear with two wheels aft and one or more wheels forward of the center of gravity. (1) Level landings-- (i) Each wheel contacting the ground simultaneously; and (ii) Aft wheels contacting the ground with forward wheels just clear of the ground. ### Page 138 Page C - 30 (2) Application of loads-- (i) Maximum design vertical loads applied alone; and (ii) The maximum design vertical lo ads applied with a drag load of at least 25 percent of the vertical load (applied at the ground contact area). (3) A 40 percent/60 percent load distri bution between wheels for configurations having two forward wheels including quadricycle. This distribution between wheels on a common axis is to be applied for the conditions of vertical loads only and for vertical loads combined with drag loads of 25 percent of the vertical loads. (4) Aircraft pitching moments are to be reacted by the forward landing gear for simultaneous wheel contact or by the angular inertia forces when the forward landing gear is clear of the ground as specified. b. Procedures . (1) The specified loading conditi ons will be used in load derivations. (2) The critical center of gravity condition will be used for each gear and gear support structure. (i) The aft center of gravity condition with the forward gear clear will normally be critical for the aft gear and gear supports. (ii) The forward cent er of gravity condition with each gear contacting the ground simultaneously will normally design forward gear elements critical for vertical loads. (iii) The forward cent er of gravity condition with the forward gear clear may result in high load factors, angular plus linear, that will greatly affect security of items of significant mass. AC 27.481 § 27.481 TAIL -DOWN LANDING CONDITIONS. a. Explanation . This section provides the criteria for tail-down landing conditions; i.e., “the maximum nose-up attitude allowing ground clearance” with ground loads acting “perpendicular to the ground.” b. Procedures . (1) The tail-down landing condition will be used to check (by analysis or test) for criticality of landing gear or support structure. This attitude generally creates the highest forward loads on the main landing gear in combination with vertical loads. ### Page 139 Page C - 31 (2) The tail-down landing condition ma y be the critical condition for both landing load factor and for energy absorption by the main gear. Section 27.725 requires that “each landing gear must be tested in the attitude simulating the landing condition that is most critical.” Where questions exist as to the critical attitude, both level landing and tail-down landing attitudes should be used in drop tests required by § 27.725. AC 27.483 § 27.483 ONE-WHEEL LANDING CONDITIONS . a. Explanation . This section gives the condition to be used for one-wheel landing conditions. Only the vertical load condition of § 27.479(b)(1) is required. b. Procedures . The one-wheel landing condition is generally critical for the landing gear-to-fuselage attachments and the landing gear elements between the attachments. Unbalanced external loads are reacted by rotorcraft inertia. AC 27.485 § 27.485 LATERAL DRIFT LANDING CONDITIONS . a. Explanation . (1) This section provides the l oading conditions which impose side (and vertical) loads on the landing gear. A level landing attitude is specified. Two main conditions required are-- (i) Only the aft wheels in contact with the ground; and (ii) All wheels contac ting the ground simultaneously. (2) Loads. The vertical loads to be appl ied with the side loads are specified as “one-half of the maximum ground reactions of § 27.479(b)(1).” These vertical loads are the level landing loads considering both contact and noncontact with the ground by the forward wheels. (i) One side load condition is specifi ed as “0.8 times the vertical reaction acting inward on one side and 0.6 times the vertical reaction acting outward on the other side” when only the aft wheels contact the ground. (ii) The other side load conditi on (for all wheels contacting the ground) specifies the 80 percent inward/60 percent outward distribution for the aft wheels and 0.8 times (80 percent) the vertical reaction for the forward wheels. b. Procedures . The loading conditions, as specified, are applied to the landing gear and attaching structure. The loads are applied at the ground contact point, except for full swiveling gear which has the load applied at the center of the axle. In other words, full swiveling gear is considered to have swiveled to a static position under the side load before the design vertical and side loads are achieved. The rotorcraft as well as the landing gear itself will be substantiated for these side load conditions. ### Page 140 Page C - 32 AC 27.493 § 27.493 BRAKED ROLL CONDITIONS . a. Explanation . This section provides two loading conditions for ground braking operations. Specific vertical loads in conjunction with drag loads (due to braking) are to be considered. The limit vertical load factor is 1.33 for condition of all wheels in contact with the ground and 1.0 for condition of aft wheels only in contact with the ground and nose wheel clear. The drag load on wheels with brakes is 0.8 times the vertical load or the drag load value based on limiting brake torque, whichever is less. The drag load value for limiting brake torque may be that determined in the performance testing to TSO C26 or equivalent, as required. b. Procedures . The braking loads are calculated from the specified criteria with the shock absorbers in their static (normal) positions and with the drag loads applied at the ground contact point. Structural substantiation of the affected structure may be accomplished by test or analysis. If tests are used, the wheel and tire assembly is commonly replaced with a test fixture so the limit loads and static deflections specified can be more accurately controlled. The test specimen should be complete enough to ensure that the landing gear structure and the attach and backup structure are adequately substantiated. AC 27.497 § 27.497 GROUND LOADING CONDITIONS: LANDING GEAR WITH TAIL WHEELS . a. Explanation . This section provides the loading conditions for landing gear designs with tail wheels. (1) Level landings are to consider the following: (i) All wheels (main and tail) c ontacting the ground simultaneously, as well as only forward main wheels contacting the ground. (ii) Maximum design vert ical loads applied alone. (iii) The maximum design vertical loads combined with a drag load of at least 25 percent of the vertical loads for both conditions. (2) Nose-up landings with only the rear wheel or wheels initially contacting the ground must be considered unless shown to be extremely remote. (3) Level landings on one forward wheel only are to be considered. Drag loads are not required. (4) Side load conditions are imposed on the main wheels and tail wheels for level landing attitudes. Criteria for full swiveling and locked tail wheels are included in this standard. ### Page 141 Page C - 33 (5) Braked roll conditions are specif ied for the level landing attitudes. (6) Rear wheel turning loads are also specified for swiveling and locked tail wheels. (7) Taxiway condition loads for the l anding gear and rotorcraft are those that “occur when the rotorcraft is taxied over the roughest ground that may reasonably be expected in normal operation.” The aircraft design load factors should not be exceeded during the evaluation. Section 27.235 contains an identical standard that applies to all types of wheel landing gear. b. Procedures . (1) The specified loading condition s are to be used in load derivations. (2) The critical center of grav ity condition is used for each gear and gear support structure. (i) The forward center of gravit y condition with the tail gear clear will normally be critical for the forward gear and gear supports. (ii) The aft center of gravity condition with the tail gear clear should be checked for criticality of security of large mass items located forward of the center of gravity. Vertical and angular accelerations are additive under this landing condition. (iii) The aft center of gravit y condition with each gear contacting the ground simultaneously will generally design tail gear elements critical for vertical loads. The other conditions are generally less severe but must be proven. (3) For tail-down landing procedures use § 27.481. The reference to “extremely remote” in § 29.497(d)(2) predates current §§ 25.1309, 29.1309, and AC 25.1309.1. This phrase has been used to require consideration of nose-up landings unless features of design are present which prevent nose-up landings or where such landings are unlikely during the life of the rotorcraft. (See paragraph AC 27.481.) (4) Use § 27.483 for one-wheel land ing procedures, paragraph AC 27.483. (5) Use § 27.485 procedures for side load conditions, paragraph AC 27.485. (6) Use § 27.493 procedures for brak ed roll conditions, paragraph AC 27.493. (7) For rear wheel turning loads, swivelin g of tail landing gears is allowed as in basic side load conditions. The side load is applied at the axle or, if the wheel is locked, the load is applied at ground contact. Rear wheels are loaded with the critical vertical static load in conjunction with an equal side load to substantiate the tail gear. ### Page 142 Page C - 34 (8) Since the rotorcraft is to be designed for load factors that will not be exceeded during taxi tests or other conditions, an instrumented taxi test program will be necessary. (Use § 27.235, paragraph AC 27.235.) AC 27.501 § 27.501 (Amendment 27- 2) GROUND LOADING CONDITIONS: **LANDING GEAR WITH SKIDS .** a. Explanation . This section provides the ground loading conditions for landing gear with skids. The loading conditions are similar to those for wheeled gear except for the following criteria which are unique to skid gears: (1) Structural yielding of elastic spri ng members under limit loads is allowed. (2) Design ultimate loads for elastic spring members need not exceed the loads obtained in a drop test with a drop height of 1.5 times the limit drop height. The rotorcraft and the landing gear attachments are subject to the prescribed design ultimate loads. (3) The gear must be in its most cr itically deflected position (similar to § 27.475). (4) Ground reactions are rationally distributed along the bottom of the skid unless otherwise specified. Section 27.501(f) concerns specific “concentrated” and arbitrary load conditions. (5) Drag loads are 50 percent of vertical reactions rather than the 25 percent for wheeled gear. (6) Side loads are 25 perc ent of the total vertical reaction rather than the 60 to 80 percent for wheeled gear. (7) Side loads are applied to one skid only (inward acting and outward acting) with resulting unbalanced moment resisted by angular acceleration. (8) A ground reaction load of 1.33 times the maximum weight is to be applied at 45° from the horizontal axis: (i) Distributed among or between the skids; (ii) Concentrated at the forward end of the straight portion of the skid tube; and (iii) Applied only to the forward end of the skid tube and its attachment to the rotorcraft. ### Page 143 Page C - 35 (9) A concentrated vertical load equal to one-half of the design limit vertical load is to be applied at a point midway between the skid tube attachments. This condition applies only to the skid tube and its attachment to the rotorcraft. b. Procedures . (1) The specified loading condition s are to be used in load derivations. (2) The critical center of gravity conditions are to be used for each gear and gear support structure. Asymmetry of the skid tubes, cross tubes, and gear attachments is to be considered in determining the critical center of gravity condition. (3) The rotorcraft and landing gear a ttachment must be substantiated for ultimate landing loads by either test or analysis utilizing an ultimate load factor of 1.5 in accordance with § 27.303. The elastic spring members may be analyzed or static tested for ultimate loads (and deflections) using either a factor of safety of 1.5 or one associated with an “ultimate” drop height of 1.5 times the limit drop height. Substantiation by “ultimate” drop tests may be used provided all combinations of critical parameters are included in the total substantiation effort. This method will require a series of tests using several test specimens or a limited number of drop tests plus further substantiations by static tests or analyses for additional critical conditions not covered by the drop test(s). AC 27.501A § 27.501 (Amendment 27- 26) GROUND LOADING CONDITIONS: **LANDING GEAR WITH SKIDS .** a. Explanation . Amendment 27-26 relaxes the previous requirements in two cases by: (1) Allowing the total sideload of § 27.501(d)(3) to be distributed “equally between skids” rather than being “applied along the length of one skid only;” and, (2) Allowing the concentrated load of § 27.501(f)(2)(ii) to be distributed over the central 33.3 percent of the skid (between skid tube attachments) rather than being “concentrated at a point midway between the skid tube attachments.” b. Procedures . The previous procedures (through Amendment 27-19) continue to apply to Amendment 27-26 except use the new load distributions. AC 27.505 § 27.505 SKI LANDING CONDITIONS . a. Explanation . This is an optional requirement for ski operations. The regulation specifies vertical loads, side loads, and torque loads (MZ) to be applied to ski installations. The four loading conditions to be applied at the pedestal bearings are: (1) Simultaneous application of Pn, up load, and Pn/4, horizontal load. ### Page 144 Page C - 36 (2) Up load of 1.33 P. (3) Side load of 0.35 Pn. (4) Torque load of 1.33 P (in foot- pounds) about the vertical axis through the centerline of the pedestal bearings. NOTE: Where P is the maximum static weight on each ski and n is the limit load factor obtained from drop tests. The load factor obtained from wheel or skid landing gear drop tests may be used. b. Procedures . Structural substantiation may be accomplished by static test or analysis using the specified loads. Skis generally have a limit load rating. The design loads derived for this standard must not exceed the rating. (TSO-C28 concerns, in part, standards for aircraft skis.) ### Page 145 Page C - 37 **SUBPART C - STRENGTH REQUIREMENTS** **WATER LOADS** AC 27.521 § 27.521 FLOAT LANDING CONDITIONS . a. Explanation . This is an optional requirement for float operations. The regulation specifies vertical loads, aft loads, and side loads to be applied to the float installations. The two loading conditions to be applied are: (1) Up-load Condition . (i) A vertical load appropriate to a landing load factor determined under § 27.473(b). (ii) A resultant water reaction pa sses vertically through the aircraft CG. (iii) An aft load equal to 25 percent of the vertical load. (2) Side-load Condition . (i) A vertical load equal to 75 percent of the vertical load for the up-load condition, equally divided among the floats. (ii) A side load at each float equal to 25 percent of the vertical load specified for the side load condition. b. Procedures . (1) The vertical load factor is det ermined by drop tests in accordance with §§ 27.473(b) and 27.725. The floats may be drop tested, or they may be assumed to have the same load factor as wheeled gear which have been drop tested. (2) Structural substantiation may be accomplished by either static tests or analysis using the specified loads. The load distribution on the floats may be realistic, based on hydrostatic pressure distributions, or conservative. ### Page 146 Page C - 38 **SUBPART C - STRENGTH REQUIREMENTS** **MAIN COMPONENT REQUIREMENTS** AC 27.547 § 27.547 (Amendmen t 27-3) MAIN ROTOR STRUCTURE. a. Explanation . This regulation requires the main rotor structure to be designed to the static load requirements of §§ 27.337 through 27.341 (vertical maneuvering loads and vertical and horizontal gust loads). In addition, the main rotor blades, hubs, and flapping hinges are specified to be designed for impact forces of each blade against its stop during ground operation and for specified limit torque at any rotational speed including zero. The torque forces (from the drive system) are distributed to the rotor blades as specified. b. Procedures . (1) Substantiation in compliance with this standard is accomplished by application of the flight loads of §§ 27.337 through 27.341 and the torque loads of § 27.361 to the rotor structure by stress analyses and/or static tests. The use of wind tunnel data as well as flight loads survey data may be used to generate and/or check the external load magnitudes and distributions. (2) Where new materials are used in the main rotor structure, such as composites containing plastics, the effects of temperature and humidity are to be considered in accordance with § 27.603, and the effects of uncertainties in manufacturing processes or inspection methods are to be considered in accordance with § 27.619. More experience is available for metallic materials, but § 27.603 requires that metallics be suitably protected against the effects of environmental conditions. (3) The design impact forces of each blade must be imposed against its stop or stops. Appropriate monitoring of the blades, hubs, flapping hinges, and stops during laboratory tests, ground endurance tests, and flight tests should ensure that the stops are sufficient for ground operation loads. The design torque loads are derived as prescribed. AC 27.549 § 27.549 (Amendment 27- 3) FUSELAGE, LANDING GEAR, AND ROTOR PYLON STRUCTURE . a. Explanation . This regulation requires that the fuselage, landing gear, and rotor pylon (including the tail fin, if any) be designed to withstand the flight loads of §§ 27.337 through 27.351, the ground loads of §§ 27.235, 27.471 through 27.497, skid loads of § 27.501, ski loads of § 27.505, water loads of § 27.521, and rotor loads of § 27.547(d) and (e). The ski and water loads pertain to optional features. Consideration is also required of -- ### Page 147 Page C - 39 (1) Auxiliary rotor thrust; (2) The torque reaction of each rotor drive system; and (3) Balancing air and inertia loads. b. Procedures . Compliance with this standard is accomplished by application of the specified aircraft loads including engine torque to the fuselage and rotor pylon structure by stress analyses and/or static tests. Drive system torque factors to be used are noted in § 27.547(e) for the main rotor structure. ### Page 148 Page C - 40 **SUBPART C - STRENGTH REQUIREMENTS** **EMERGENCY LANDING CONDITIONS** **AC 27.561 EMERGENCY LANDING CONDITIONS--GENERAL.** a. Explanation . (1) Occupant protection . The occupants should be protected as prescribed from serious injury during an emergency/minor crash landing on water or land for the conditions prescribed in the standard. The standard states that each occupant should be given every reasonable chance of escaping serious injury in a minor crash landing. In addition, the occupants shall be protected from items of mass inside the cabin as well as outside the cabin. For example, a cabin fire extinguisher must be restrained for the load factors prescribed in this section (reference § 27.1411(b)(2).) A transmission or engine must be restrained to the load factors in § 27.561(b)(3) if located above or behind the occupants. (2) Load factor determination . The standard in § 27.561(b)(3) specifies certain ultimate inertia load factors but allows a lesser downward vertical load factor by virtue of a 5 FPS ultimate rate of descent. (3) Retractable landing gear . For rotorcraft equipped with retractable landing gear only the retracted configuration must be considered. (4) Door and exit protection . The minor crash conditions contained in § 27.561(b)(3) should also be considered in designing doors and exits (§ 27.807(b)(4)). (5) External load considerations . The load factors of § 27.561 and the criteria of § 27.562 are not directly applicable to external load systems. This is because in emergency crash scenarios that involve external loads, the external load is neither typically subjected to the same minor crash loads (§ 27.561) as is the rotorcraft hull and its internal occupants nor are all of the occupant protection criteria (§ 27.562) needed or practicable to apply. Appropriate safety for external load carriage systems is provided by the criteria of § 27.865. Safety standards for external load attaching means are provided in § 27.865. b. Procedures . (1) The design criteria report or another similar report of the rotorcraft structural limits should contain the (ultimate) minor crash condition load factors. (2) Section 27.785 (paragraph AC 27.785) concerns application of this design standard to seats, berths, belts, and harnesses. ### Page 149 Page C - 41 (3) The ultimate design landing and maneuvering load factors may exceed the minor crash condition load factors. The highest load factor derived shall be used. (i) For example, for light weight conditions, the ultimate maneuvering load factor may be 5.25g as specified in § 27.337. (ii) The ultimate vertical l anding load factors derived from §§ 27.471 through 27.521, whichever is appropriate for the design, may exceed the 4.0g down load factor in this section. The rotorcraft landing case design limit contact velocity shall range from 6.5 to 8.3 FPS (see §§ 27.473 and 27.725). (4) As specified in (b)(3)(iv) of the standard, the downward load factor is 4.0 or a lower design load factor may be used. (i) The lower load factor relates to a rotorcraft impacting a flat, hard landing surface at 5 FPS (ultimate) vertical rate of descent. The load factor derived for each unique design is a function of the aircraft impact/crushing characteristics. (ii) The 4.0g down load factor ca se is related to either a fixed or retractable gear rotorcraft. This condition is not dependent on impact characteristics of the rotorcraft. (iii) As noted in paragraph b(3) abov e, the design landing load factors may exceed each of the two previous cases and would then become the prominent design (vertical load) parameter for seats, transmissions, fire extinguishers, etc. (5) Items of mass such as fire ex tinguishers, nav-com equipment, life rafts, engines, and/or transmissions shall be restrained for the appropriate load factors. (6) Cargo/baggage compar tments separated from the passenger compartment shall be designed for load factors specified in § 27.787. The conditions in § 27.561 are excepted from that standard. AC 27.561A. § 27.561 (Amendment 27- 25) EMERGENCY LANDING CONDITIONS -GENERAL. a. Explanation . Amendment 27-25 adds or increases the design static load factor of § 27.561 in two areas. The addition of these load factors eliminates the 5 FPS ultimate descent velocity criteria of unamended § 27.561(b)(3). (1) The design static load factors for the cabin in § 27.561(b)(3) are increased in concert with the dynamic test requirements of new § 27.562. (2) Design static load factors are added in § 27.561(c) for external items of mass located above and/or behind the crew and passenger compartment. ### Page 150 Page C - 42 b. Procedures . The procedures of paragraph AC 27.561 continue to apply except the new load factors of § 27.561 should be used. Penetration of any items of mass into the cabin or occupied areas should be prevented. AC 27.561B. § 27.561 (Amendm ent 27-30) EMERGENCY LANDING CONDITIONS - GENERAL . a. Explanation . (1) Amendment 27-30 adds paragraph (d) wh ich lists specific load factors for the fuselage structure in the area of internal fuel tanks located below the passenger floor level. For other locations, the fuselage structure is to be designed to resist crash impact loads prescribed in § 27.561(b)(3) for fuel tanks located within the cabin area; or § 27.561(c) for fuel tanks located behind or above the occupant area. These load factors are provided to prevent crash induced fuel tank ballistic hazards to occupants and to also protect the fuel tank from rupture as prescribed. The landing gear must be retracted if the rotorcraft is equipped with retractable gear. (2) Section 27.952(b) provides specific load factors for the fuel tanks which are identical to the load factors stated in § 27.561. Paragraph AC 27.952 provides information and guidance for § 27.952 and may be used in conjunction with this paragraph. b. Procedures . The procedures of paragraphs AC 27.561 and AC 27.561A continue to apply except new load factors are established for fuel tanks located below the passenger floor level. Each fuel tank and its installation are subject to the loads stated in the standard. The load factors are determined by the fuel tank location. (1) The crash impact load factors for the airframe structure surrounding the underfloor fuel tanks are specified in § 27.561(d). The fuselage structure must be designed to resist the specified crash impact loads and to help protect the fuel tank from rupture. If equipped with retractable landing gear, the effects of the landing gear on fuel system rupture should be considered in both the retracted and unretracted configurations for fuel cell hazard purposes, only, in accordance with AC 27.952. (2) Section 27.952(b) (see AC 27.952) specifies the design load factors for crash resistant fuel systems in an otherwise survivable impact. This section relates to § 27.561(d) as follows. The § 27.952 load factors are for the fuel tanks, other significant mass items in the fuel system, and their attachment to the rotorcraft airframe for both occupant survivability and retention of fuel in a survivable impact; whereas, the § 27.561(d) load factors only apply to the rotorcraft airframe surrounding the underfloor fuel tanks and their installation for the same reasons. These two sets of load factors are not additive. They are applied separately (as design ultimate load factors) to the portions of the rotorcraft to which they are specified to apply. The application of the § 27.561(d) load factors is described as follows. The loads generated by § 27.561(d) are intended to be applied to the airframe structure surrounding the fuel cell to ensure ### Page 151 Page C - 43 that the entire airframe structure provides the appropriate level of crash resistance (i.e., stiffness, crushability, crushing rate, energy absorption capability, etc.) and to ensure that the airframe structure’s failure modes (e.g., buckling, creation of sharp edges, structural spears, etc.) are such that fuel cell rupture (and the resultant post crash fire potential) is mitigated to the maximum practicable extent in a otherwise survivable emergency landing. Each fuel cell (and major fuel cell component) creates an applied load on the airframe in an emergency landing condition. These loads are determined by multiplying the worst case mass of the fuel cell (i.e., a full fuel cell) by the load factors of § 27.561(d). These loads are then applied (utilizing the appropriate design load paths) to the airframe structure surrounding the fuel cell to help design the structure for optimal crash resistance. Added stiffness effects for both a full and less than full fuel cell should be considered in the design process. A significantly less than full fuel cell will typically not have any significant stiffness effects, since in a less than full condition the fuel cell cannot typically transfer load hydraulically. AC 27.561C. § 27.561 (Amendm ent 27-32) EMERGENCY LANDING CONDITIONS - GENERAL . a. Explanation . Amendment 27-32 adds a new rearward emergency load factor of 1.5g to both §§ 27.561(b)(3)(v) and 27.561(c)(5). The addition of the 1.5g rearward load factor in § 27.561(b)(3)(v) is to provide an aft ultimate load condition for substantiation of the restraints required for retention of both occupants and significant items of mass inside the cabin that could otherwise come loose and cause injuries in an emergency landing. The addition of the 1.5g rearward load factor to § 27.561(c)(5) is to provide an aft ultimate load condition for substantiation of the support structure for retention of significant items of mass above and forward of the occupied volume(s) of the rotorcraft that could otherwise come loose and injure an occupant in an emergency landing. Amendment 27-32 also increases the forward, sideward, and downward emergency load factors of § 27.561(c)(2), (c)(3), and (c)(4), respectively; for retention of items of mass above and behind the occupied volume(s) that could otherwise come loose and injure an occupant in an emergency landing. b. Procedures . The procedures of paragraphs AC 27.561, AC 27.561A, and AC 27.561B continue to apply except the newly specified load factors must be used. A list of the significant items of mass to be considered should be compiled by the applicant and approved by the certifying authority. AC 27.562 § 27.562 (Amendment 27-25) EMERGENCY LANDING DYNAMIC CONDITIONS. a. Explanation . Amendment 27-25 adds new requirements for the dynamic testing of all seats in rotorcraft. This paragraph is rewritten to incorporate the guidance previously documented in AC 20-137 dated 3/30/92. b. Background . Improved occupant restraint in civil rotorcraft is addressed in Amendments 27-25 to the airworthiness standards, which add two dynamic crash ### Page 152 AC 27-1B, Chg 1 2/12/03 Page C - 44 impact design conditions for seat and occupant restraint systems. This amendment also prescribes a shoulder harness for each occupant and adopts human impact injury criteria as a measure for occupant protection for the dynamic crash impact conditions. In addition, these amendments significantly improve occupant protection for normal category rotorcraft in a survivable emergency landing. This advisory material addresses the dynamic test conditions and the related pass-fail injury criteria for the dynamic test conditions. This material pertains to single as well as multiple seats and tandem arrangements of the seats in rotorcraft. (1) Dynamic test methods . This guidance focuses on the use of dynamic test methods for evaluating the performance of rotorcraft seats, restraint systems, and certain related interior systems for demonstrating structural strength and the ability of those systems to protect an occupant from possible injuries in an emergency landing environment represented by the standard. These test methods differ from static test methods, which are limited to demonstrating only the structural strength of the seat or restraint system under ultimate load for at least 3 seconds. This guidance contains sources for appropriate test procedures and provides some insight into the logic of these procedures. It also defines, in part, test facility and equipment characteristics necessary for conducting these tests. (2) Standardized test methods . Dynamic tests are often conducted at a specially equipped facility, one other than that owned by the designer or manufacturer of the test article. To obtain consistent test results, it is necessary to specify the critical test procedures in detail in the test plan, and then carefully follow these procedures when conducting the tests. This guidance defines certain critical procedures for accomplishing the tests of the seat and restraint systems and assessing the data obtained in the tests. Many of these procedures are accepted as standards by government and commercial test facilities and have been modified in this guidance only as necessary for the specific testing of rotorcraft systems. (3) Relationship of dynam ic tests to design standards. This guidance describes test procedures useful in assessing the performance of a rotorcraft seat, restraint system, and interior system. However, it is impractical to conduct sufficient tests for assessing the performance of the system throughout its entire range of possible uses in unique interior arrangements. The seat, restraint system, and related interior system should be designed for the range of occupants and environments for which it is expected to perform, not just for the dynamic test conditions described in this guidance. (i) Occupant size . The dynamic tests are conducted with a specific, acceptable, standard anthropomorphic test dummy (ATD) representing a 50th percentile male occupant. Energy absorbing systems, restraint system loads and anchorage locations, seat adjustments, seat pitch (for multiple seat rows), head strike envelopes, etc., are typical factors directly influenced by occupant size. (ii) Seat position and location . The tests should be sufficient to represent the range of performance expected of a seat and restraint system. A seat, especially an ### Page 153 Page C - 45 adjustable flight crew seat, should be qualified for those positions approved for take-off and landing. As with static test procedures the seat is also tested to the most critical condition for the dynamic tests. For an adjustable flight crew seat, as an example, the full-up position and longitudinal impact case are expected to be the critical condition. But these dynamic tests and occupant injury assessment provide a systems approach to qualification. It is therefore necessary to test adjustable seats at the design position for the ATD. Two tests would be required to demonstrate compliance with the strength standards and with the occupant injury criteria. Alternatively adjusting the flight crew seat to its highest position with the interior features, such as an instrument panel shield, raised to maintain the proper perspective or relation to the ATD, is considered an acceptable test procedure for demonstrating compliance with the structural and occupant injury requirements for the seat and its location in a particular cockpit arrangement. (iii) Test conditions . Only two minimum impact tests are described in the dynamic test procedures discussed in this guidance. These procedures address the tests needed to demonstrate compliance for a typical seat and restraint system installation. Additional tests may be necessary to demonstrate compliance for other types or variations of seat and restraint system installations. For example, while only one lateral load direction is specified in the tests, the system should perform properly when similarly loaded from either side. (iv) Floor deformation . The test procedures require evaluating the effect of certain sidewall or floor deformation. The seat and its attachments and the restraint system should also perform properly if no floor deformation is present. (v) Head impact . Should such contact occur, head impact with a seat back or the interior of the rotorcraft is evaluated by using a Head Injury Criterion (HIC), which can be measured directly in the tests discussed in this guidance or in supplementary tests of the interior. The design of the interior should protect the head from serious injury throughout the head strike envelope, not just along the head strike paths demonstrated in the test conditions discussed in this guidance. (vi) Emergency egress . Standards for emergency evacuation of the rotorcraft are contained in FAR Part 27. The objective is to allow each occupant to leave the seat and rapidly evacuate the rotorcraft using an exit on either side of the rotorcraft. c. Dynamic Test Methods and Facilities . (1) General . A minimum of two dynamic tests are used to assess the performance of the rotorcraft seat, restraint system, and related interior system. The seat, the restraint, and the nearby interior all function together as a system to protect the occupant during emergency landing. The specific test conditions are shown in Figure AC 27.562-1. Explanations of the test conditions are as follows: ### Page 154 AC 27-1B, Chg 1 2/12/03 Page C - 46 (i) Test 1 . The test determines the protection provided when the impact environment is such that the resulting predominant impact load component (vertical) is directed along the spinal column of the occupant in combination with a horizontal (longitudinal) component. Protection against spinal injury is important and it may be necessary to provide energy absorbing (load limiting) or attenuation capability in the seat system in order to comply with the human injury criteria specified in § 27.562 (c)(7). (ii) Test 2 . The test determines the protection provided in an impact where the predominant impact load component is in the longitudinal direction in combination with a lateral component. Evaluation of head injury protection is important in this test if the head could strike some interior portion of the rotorcraft or a forward seat. Chest or spinal column injury, which might result from the upper torso restraint, is also evaluated in this test. (iii) Tests 1 and 2 . These test conditions are also significant for the structural strength of the system. Both tests should be used to assess submarining (where the seat belt slips above the ATD pelvis) and rollout of the upper torso restraint system particularly with single, diagonal torso restraint belts. Since external crash forces frequently cause significant structural deformation, simulated floor deformation is specified for the tests to prove the seat design can accommodate the relative deformation between the seat and the floor or sidewall and still function without imposing excessive loads on the seat, the attachment fittings, or floor tracks. ### Page 155 (OCR) a forward facing Illustration shows Test 1 Test 2 seat * ATD Inertial load shown by atrow: 10° Min. Vi. fps 30 12 Max. tr. sec. _ 031 . 071 Min. G 30 18.4 Deform floor: degrees roll 10 10 degrees pitch 10 10 deceleration tr = rise time V1 = Impact velocity FIGURE AC 27.562-1. Seat Restraint System Dynamic Tests Normal and Transport Category Rotorcraft Page C - 47 ### Page 156 AC 27-1B, Chg 1 2/12/03 Page C - 48 (2) Test facilities . A test proposal is prepared for certification authority approval and should reflect the capability of the facility. It should be noted that a number of test facilities could be used to accomplish dynamic testing. Test facilities can be grouped into categories based on the method they use to generate the impact pulse (i.e., accelerators, decelerators, or impact with rebound) and whether the facility is a horizontal (sled) design or a vertical (drop tower) arrangement. As in all certification compliance tests, a test proposal, which may refer to certain specific or generic test equipment, is approved prior to testing. The test may be conducted anywhere, within certain availability or mutually convenient constraints, as long as the test is conducted in accordance with the approved test plan and properly witnessed. (i) Facility Characteristics or Features . Each of the facilities has characteristics that may have advantages or disadvantages with regard to the dynamic tests discussed in this guidance. One concern is the rapid sequence of acceleration and deceleration that must take place in the tests. In a landing impact, the acceleration phase (flight) is gradual and usually well separated in time from the deceleration (crash impact) phase. In a test, the deceleration usually closely follows the acceleration. When assessing the use of a facility for the specific test procedures outlined in the recommendations, it is necessary to assess the possible consequences of this rapid sequence of acceleration and deceleration on the test articles and ATD. The standard accommodates the different facilities that are or may be available for the applicant’s use. That is, the standards dictate the peak acceleration with a tolerance as stated in this AC. The “decay” in deceleration with respect to time is not dictated, thereby allowing for the different test facility equipment characteristics. (A) Deceleration sled facilities . In an aircraft crash, the impact takes place as a deceleration, so loads are applied more naturally in test facilities that create the test impact pulse as a deceleration. Since it is simpler to design test facilities to extract energy in a controlled manner than to impart energy in a controlled manner, several different deceleration sled facilities can be found. The deceleration sled facility at the FAA’s Civil Aeromedical Institute (CAMI) was referred to in developing the test procedures discussed in this and similar AC’s related to airplanes. ( 1 ) The Acceleration Phase. Sufficient velocity for the test impact pulse acquired in this phase can distort the test results if the acceleration is so high that the test articles or ATD are moved from their intended pre-test position. This inability to control the initial or onset conditions of the test would directly affect the test results. This can be avoided by using a lower acceleration for a relatively long duration and by providing a coast phase (in which the acceleration or deceleration is nearly zero) prior to the impact. This allows any dynamic oscillation in the test articles or the ATD that might be caused by the acceleration to decay. To guard against errors in data caused by pre-impact accelerations, data from the electronic test measurements (accelerations, loads) should be reviewed for the time period just before the test impact pulse to make sure all measurements are at the baseline (zero) level. Photometry film taken of the ### Page 157 Page C - 49 test should also be reviewed to make certain that the ATD’s used in the test and the test articles were all in their proper position prior to the test impact pulse. ( 2 ) Orientation of test article. The horizontal test facility readily accommodates forward-facing seats in both tests discussed in this guidance, but problems can exist in positioning the test ATD in Test 1 if the seat is a rear or sidefacing seat. In these cases, the ATD’s tend to fall out of the seat due to the force of gravity and must be restrained in place using breakaway tape, cords, or strings. Since each installation will present its own problems, there is no simple, generally applicable, guidance. Attention should be given to positioning the ATD against the seat back and to proper positioning of the ATD’s arms and legs. It will probably be necessary to build special supports for a breakaway restraint so that the restraint will not interfere with the function of the seat and occupant restraint system during the test. Photos of the test from “side of track cameras” should be reviewed to make sure that the breakaway restraint did break (or become slack) in a manner that did not unduly influence the motion of the ATD or the test articles during the test. (B) Acceleration sled facilities . Acceleration sled facilities, usually based on the Hydraulically Controlled Gas Energized (HYGE) accelerator device, provide the impact test pulse as a controlled acceleration at the beginning of the test. The test item and the ATD are installed facing in the opposite direction from the velocity vector, opposite from the direction used on a deceleration facility, to account for the change in direction of the impact. There should be no problem with the ATD or the test items being out of position due to pre-impact sled acceleration, since there is no sled movement prior to the impact test pulse. Because of this characteristic the applicant may prefer this type of a facility. ( 1 ) Test pulse. After the impact test pul se, when the sled is moving at the maximum test velocity, stop the sled safely. Most of the facilities of this design have limited track length available for deceleration, so that the deceleration levels can be relatively high and deceleration may begin immediately after the impact test pulse. Since the maximum response of the system usually follows (in time) the impact test pulse, any sled deceleration, which takes place during that response will affect the response and change the test results. The magnitude of change depends on the system being tested, so that no general “correction factor” can be specified. The effect can be minimized if the sled is allowed to coast, without significant deceleration, until the response is complete. ( 2 ) Test results. If the seat or re straint system experiences a structural failure during the test pulse, the post impact deceleration can increase the damage and perhaps result in failures of unrelated components. This will complicate the determination of the initial failure mode and make product improvement more difficult. One other consideration is that the photometry film coverage of the response to impact test pulse must be accomplished when the sled is moving at near maximum velocity. Onboard cameras or a series of trackside cameras are usually used to provide film coverage of the test. Since onboard cameras frequently use a wide-angle lens placed ### Page 158 AC 27-1B, Chg 1 2/12/03 Page C - 50 close to the test items, it is necessary to account for the effects of distortion and parallax when analyzing the film. The acceleration sled facility faces the same problems in accommodating rearward-facing or side-facing seats in Test 1 as the deceleration sled facility, and the corrective action is the same for both facilities. (C) Impact-with-rebound sled facilities . One other type of horizontal test facility used is the “impact-with-rebound” sled facility. On this facility, the impact takes place as the moving sled contacts a braking system, which stores the energy of the impact, and then returns the stored energy back to the sled, causing it to rebound in the opposite direction. This facility has an advantage over acceleration or deceleration facilities in that only one-half of the required velocity for the impact would need to be generated by the facility (assuming 100 percent efficiency). Thus the track length can be shortened, and the method of generating velocity is simplified. The disadvantages of this facility combine the problems mentioned for both acceleration facilities and deceleration facilities. Since one of the reasons for this type of facility is to allow short track length to be used, it may be difficult to obtain sufficiently low acceleration just before or after the impact pulse to resolve data error problems caused by significant pre-impact and post-impact accelerations. (D) Drop towers . Vertical test facilities can include both drop towers (decelerators) and vertical accelerators. Vertical accelerators, which can produce a longer duration/displacement impact pulse, may not be available. However, drop towers are one of the easiest facilities to build and operate and are frequently used. ( 1 ) Acceleration phase. In these facilit ies, the pull of earth’s gravity is used to accelerate the sled or guided test fixture and test article to specified impact velocity to avoid the use of a complex mechanical accelerating system. Reproducing the required impact pulse may require extensive development tests for the facility. Unfortunately, these facilities are more difficult to use for conducting Test 2, particularly for typical forward-facing seats. ( 2 ) Test article. In preparing for (longi tudinal) Test 2, the seat should be installed at an angle according to the standards such that the ATD’s tend to fall from the seat due to gravity. The restraint system being tested cannot hold the ATD against the seat unless tightened excessively and will not usually locate the head, arms, or legs in their proper position relative to the seat. Design and fabrication of an auxiliary “break-away” ATD positioning restraint system just for this test are usually a complex task. The auxiliary restraint should not only position the ATD against the seat (including maintaining proper seat cushion deflection) during the pre-release condition of 1 g, it should also maintain the ATD in that proper position during the free fall to impact velocity when the system is exposed to zero g, and then it should release the ATD in a manner that does not interfere with the ATD response to impact. The usual sequence of 1 g/0 g impact, without the possibility of a useful “coast” phase, as done in horizontal facilities, causes shifts in initial conditions for the test impact pulse that can affect the response to the impact. The significance of this undesired movement will depend on ### Page 159 Page C - 51 the dynamic characteristics of the system under test, and these characteristics are seldom known with sufficient accuracy to achieve the response initially. ( 3 ) Other facets. In addition, the earth’s gravity will oppose the final rebound of the ATD into the seat back, so that an adequate test of seat back strength and support for the ATD cannot be obtained. The problems in Test 1, or with rear-facing seats in Test 2, are not as difficult because the seat will support the ATD prior to the free fall. However, the zero g condition free fall that exists prior to impact will allow the ATD to “float” in the seat restraint system, perhaps changing position and certainly changing the initial impact conditions if movement occurs. Again, the development of a satisfactory auxiliary breakaway restraint system to assure correct pre-impact condition is difficult. (ii) Test Fixtures. A test fixtur e is usually required to position the test article on the sled or drop carriage of the test facility and to represent the aircraft’s structure floor, sidewall, bulkhead, etc. It holds the attachment fittings or floor tracks for the seat, provides the floor and sidewall deformation needed for the test, and provides a floor or footrest for the ATD, and it positions the pertinent interior items, such as instrument panels, sidewalls, bulkheads, a second row of seats, if required, for successful performance of the tests, and otherwise simulates the rotorcraft for the test. The text fixture is usually fabricated of heavy structural steel and does not necessarily simulate lightweight aircraft design or construction. The details of the fixture will depend upon the requirements of the test articles, but provisions for the specified floor and sidewall deformation are needed. (A) Purpose of floor or sidewall deformation . The purpose of using pitch and roll deformation for the tests is to demonstrate that the seat/restraint system will remain attached to the airframe and perform properly for the tests, although the structure and seat may be more severely deformed by the forces associated with a particular crash. Typical design deficiencies addressed by the test conditions include, but are not limited to, the following: ( 1 ) Concentrated loads may be imposed on floor fittings (studs) or tracks by seat leg attachment fittings which fit tightly or are clamped to a track or fitting, and which do not have some form of relief (especially lateral roll relief) incorporated in the design. These joint fittings can concentrate the forces on one lip of the floor and sidewall track or stud and may break the joint (track or the fitting). ( 2 ) Similarly, loads can be concentrated on one edge of a floor track or stud fitting having an “I,” “bulb head” or “mushroom” cross section and may prematurely break the flange or the fitting. ( 3 ) Detents, pins, or collars which lock t he seat leg fitting to the floor track can become disengaged, or the mechanism which is used to disengage the detents, pins, or “dogs” can be actuated and release the seat as the seat or airframe deforms. ### Page 160 AC 27-1B, Chg 1 2/12/03 Page C - 52 ( 4 ) Seat assemblies that provide an energy absorbing system between a seat “bucket or pan” and a seat frame attached to the floor may not perform properly for a pre-loaded seat frame attached to the floor or sidewall. Deformation of the seat frame may cause the energy absorbers to receive unanticipated loads or cause excessive friction in the guides between the seat bucket and seat frame to lock the energy absorber in place. ( 5 ) Restraint system anchorages attached to the airframe structure may be significantly displaced relative to the seat if the seat deforms during the test, and that displacement may inhibit proper performance of the seat/restraint system. This is especially critical for the necessary vertical stroking or displacement. (B) Floor Deformation . The pitch and roll displacement is intended to evaluate the track or stud and leg fitting joint (axis) tolerance to angular misalignment and not necessarily axis translational displacement. ( 1 ) For the typical aircraft seat. For a multiple or single person seat, with four seat legs mounted in the airframe on two parallel tracks, the floor deformation test fixture may consist of two parallel beams, a “pitch beam” which pivots about a lateral (y) axis, and a “roll beam” which pivots about a longitudinal (x) axis. The beams can be made of any fairly rigid structural form, box, I-beam, channel, or other appropriate cross section. The pitch beam should be capable of rotating in the x-z plane up to +/- 10° relative to the longitudinal (x) axis. The roll beam should be capable of a +/- 10° roll about the axis of the seat attachment/fitting joint (centerline of floor track or fittings). (See Figure 27.562-2 for a schematic of an installation.) A means should be provided to fasten the beams in the deformed positions. ( 2 ) Seat and floor interface. The beams should have provision for installing floor tracks or other attachment fittings on their upper surface in a manner that does not alter the above-floor strength of the track or fitting. The track or other attachment fittings should be representative (in above-floor configuration shape and strength) of those used in the rotorcraft. Structural elements below the surface of the floor that are not considered part of the floor track or fitting may be omitted in the installation. The seat having four legs should then be installed on the beams so that the rear seat leg attachment point is near the pitch beam axis of rotation, and the seat positioning pins or locks are fastened in the same manner as specified in the test proposal and as would be used in the rotorcraft, including the adjustment of “anti-rattle” mechanisms, if employed. ( 3 ) Test set-up. The remainder of the test preparations would then be completed (ATD installation and positioning, instrumentation installation, adjustment and calibration, camera checks, etc.). The “floor deformation” would be induced as the final action before the test is accomplished. The roll beam should first be rotated 10° and locked in place, and then the pitch beam should be rotated 10° and locked in place. The direction of rotation would be selected to produce the most critical loading condition on the seat and floor track or fitting. If the seat is fairly flexible, it may be possible to ### Page 161 Page C - 53 rotate the beams by manual effort, perhaps using removable pry bars to gain mechanical advantage. However, rotation of the beams used for testing a stiff seat frame is likely to require greater effort than can be accomplished manually, and the use of removable hydraulic jacks or other devices may be necessary. If this condition is expected, provision should be made for appropriate loading points when designing the fixture. This condition is most likely to be encountered when rotating the pitch beam. The test facility personnel should adhere to appropriate safety provisions during the deformation process. The test fixture may be designed to adjust to fit a wide range of seat designs, including leg spacing, that may be encountered. ### Page 162 (OCR) AC 27-1B, Chg 1 2/12/03 **ORIGINAL** POSITION NOTE: **FOR CLARITY,** **LOAD TRANS-** **DUCERS ARE** **NOT SHOWN POSITION** **AFTER FLOOR** **DEFORMATION** **PITCH AXIS** **UNDER REAR** **LEG ATTACHMENT** ORIGINAL ORIGINAL: POSITION POSITION _ROLL BEAM > PITCH 10° PITCH BEAM 10° ROLL TO BE TESTED . MINIMUM SEAI TRACK **ROLL LEFT OR RIGHT** **DETAIL (A)** **EXAMPLE OF AIRCRAFT** TRACK BEAM FIGURE AC 27.562-2 - Schematic Floor Deformation Fixture; Seat Legs Attached at Floor Level Page C - 54 ### Page 163 Page C - 55 (C) Alternative configurations . The preceding discussion described the fixture and floor deformation procedure that would be used for a typical seat that has four seat legs and four attachments to the fuselage floor. These test procedures may be adapted to seats having other designs. Special test fixtures may be necessary for different configurations. The following methods, while not covering all possible seat designs, provide guidance for the more common configurations of seats: ( 1 ) Rotorcraft seats with three legs ma y have one central leg in front or back of the seat and one leg on each side of the seat. The central leg should be held in its undeformed position as pitch deformation is applied to one side leg and roll to the other. ( 2 ) Seats that are “integr al” with the structure without floor or sidewall attachment devices and with continuous attachments such as rows of rivets or screw, etc., are excluded from the deformation, misalignment, or preload prior to test impact. Similarly bulkhead-mounted seats, solely mounted to a bulkhead, are excluded from the deformation requirement. The test fixture could represent the seat and structure or a rigid bulkhead or an actual bulkhead panel. If a rigid bulkhead installation is used, the test fixture should transfer loads to the seat restraint system through components equivalent to the seat attachment fittings and surrounding bulkhead panel, which exist in the actual installation. Similar guidelines apply to integral seats. ( 3 ) Seats that are attached to both the floor and a bulkhead would be tested on a fixture that positions the bulkhead surface in a plane through the axis of rotation of the pitch beam. The bulkhead surface should be located perpendicular to the plane of the floor (the rotorcraft floor surface, if one were present) in the undeformed condition or in a manner appropriate to the intended installation. Either a rigid bulkhead simulation or replica or an actual bulkhead panel may be used. If a rigid bulkhead simulation is used, the test fixture should transfer loads to the seat restraint system through components equivalent to the attachment fitting and surrounding bulkhead panel that would exist in the actual installation. The seats would be attached to the bulkhead and the floor in a manner representative of the rotorcraft installation, and the floor, as represented in the test, would then be deformed as described in paragraph b(2)(ii)(B). ( 4 ) Seats mounted between fuselage sidewa lls or to the sidewall and floor of an airplane should be tested in a manner simulating rotorcraft fuselage cross-section deformation (e.g., from circular or rectangular to flattened circular or rectangular or ellipsoidal shape) during a severe impact. The 10° roll would simulate the change in fuselage shape. Brackets should be fabricated to attach the seat to the sidewall test fixture at the same level above the fixture “floor” that would represent the installation above the rotorcraft floor. The sidewall bracket or rail should be located on the “roll” beam. It is envisaged that the sidewall rolls outward 10° about an axis at the floor and sidewall juncture. Then, as the beams are rotated to produce the critical loading condition, the combined angular and translational deformation would simulate the ### Page 164 AC 27-1B, Chg 1 2/12/03 Page C - 56 deformation at the sidewall attachment during a landing impact. (See Figure 27.562-3 for a schematic of an installation.) ( 5 ) Seats that are cant ilevered from one sidewall without connection to another structure would not be subject to floor deformation. However, sidewall deformation is likely, and should be considered by warping the entire sidewall attachment plane, or the attachment points of the seat, 10° to represent the most likely fuselage sidewall deformation. This is intended to evaluate a critical condition for seat attachment or seat and occupant restraint system performance. Either a rigid sidewall simulation or an actual sidewall panel may be used. If a rigid sidewall simulation is used, the test fixture should transfer loads to the seat through components equivalent to the attachment fitting as well as the surrounding sidewall to replicate the actual installation. ( 6 ) Side-facing seats, occupiable for takeoff and landing, are subject to the specified dynamic test conditions. Compliance with the structural requirements should be demonstrated for side-facing seats using the same conditions for the test and pass/fail criteria as for fore- and aft-facing seats. The seat should be loaded in the most critical case structurally. Means of restraining the ATDs may need to be adapted to ensure adequate retention during the test. The application of floor distortion will need to be assessed on an individual basis, depending on the design and the method of attaching the seat. ( 7 ) A seat assembly for multiple occupants may have more than two pairs of legs. If the assembly uses a uniform cross section, deformation of only the outer leg assemblies is sufficient. The inner leg pairs may be maintained in the normal, undeformed position for the dynamic tests. (D) Multiple Row Test Fixtures . In tests of passenger seats normally installed in rows in a rotorcraft, head impact conditions should be evaluated by tests using at least two rows of seats. This allows direct measurements of the head injury data if secondary head impact occurs and demonstrates the effect of the interaction loads between rows; e.g., due to occupant contact with the front row. (That is, ATD leg contact does not overload the front row.) These conditions are usually critical only on Test 2. The single seat row fixture used for the test should be used to position the front (first) seat row and provide appropriate floor deformation to that row. The test is critical for the first row strength. An additional simple fixture may position the second seat row in the proper location and need not provide floor deformation. The second row should be fully occupied unless it is not as critical a condition for the first seat row. Representative seat cushions and torso restraint systems should be used on both seat rows. The allowable seat pitch (longitudinal spacing) can be determined by analysis of previous test data or limited by type design data and information for the most critical condition for head or leg impact against relatively stiff structure in the first seat row. Operational limitations that specify the allowable seat pitch of the seats in rotorcraft may be considered also. No impact surface such as seats, bulkheads, etc., may be needed ### Page 165 Page C - 57 for the ATD in the first seat row unless such a surface is within the expected head strike envelope whenever the seats are installed in rotorcraft. ### Page 166 (OCR) AC 27-1B, Chg 1 2/12/03 **ORIGINAL** POSITION NOTE: POSITION **FOR CLARITY, AFTER FLOOR** **LOAD TRANSDUCERS DEFORMATION** **ARE NOT SHOWN REPRESENTATION** **SEAT TO SIDEWALL** **ATTACHMENT -RIGID TEST** **STRUCTURE** **PITCH AXIS ORIGINAL** **UNDER REAR POSITION** LEG ATTACHMENT ORIGINAL UP OR DOWN POSITION 10° PITCH PITCH BEAM ROLL AXIS - OUTBOARD 10° ROLL ......... FIGURE AC 27.562-3 - Schematic Test Fixture; Side Wall Mounted Seat Page C - 58 ### Page 167 Page C - 59 (E) Other fixture applications . Test fixtures should provide a flat footrest for an ATD used in tests of passenger and attendant seats. Flightcrew seats associated with special foot rests or foot-operated controls may use simulated footrests. The surface of the footrest should be covered with carpet (or other appropriate material) and be at a position representative of the undeformed floor or control. Test fixtures may also be necessary to provide guides or anchors for torso restraint systems or for holding instrument panels or bulkheads if necessary for the proposed tests. If these provisions are necessary, the installation should represent the configuration of the installation and be of adequate structural strength to withstand the expected test loads. (iii) Instrumentation . Electronic and photographic instrumentation systems are essential to properly record the information for the tests discussed in this AC. Electronic instrumentation is used to measure accelerations and forces required for verifying the test environment and for measuring most of the pass/fail criteria and the floor (seat) attach loads. Photographic instrumentation is used for recording the overall qualitative results of the tests, for confirming that the lap safety belt remained on the ATD’s pelvis (no submarining), and that the upper torso restraint straps remained on the ATD’s shoulder, and for recording the relative deformation of the seats as it may influence rapid evacuation of the rotorcraft by the occupants. Paragraph d.(10), of this guidance contains allowable seat deformation information related to an aisle, passageway, access to exits, and so forth. (A) Electronic instrumentation . Electronic instrumentation should be accomplished in accordance with the Society of Automotive Engineers Recommended Practice SAE J211, Instrumentation for Impact Tests. In this practice, a data channel is considered to include all of the instrumentation components from the transducer through the final data measurement, including connecting cables and any analytical procedures that could alter the magnitude or frequency content of the data. Each dynamic data channel is assigned a nominal channel “class” equivalent to the high frequency limit for that channel, based on a constant output/input ratio vs. frequency response plot which begins at 0.1 Hz (+1/2 to –1/2 db) and extends to the high frequency limit (+1/2 to – 1 db). Frequency response characteristics beyond this high frequency limit are also specified. When digitizing data, the sample rate should be at least five times the –3 db cutoff frequency of the pre-sample analog filters. Since most facilities set all pre-sample analog filters for Channel Class 1,000 and since the –3 db cutoff frequency for Channel Class 1,000 is 1,650 Hz, the minimum digital sampling rate would be about 8,000 samples per second. For the dynamic tests discussed in this guidance, the dynamic data channels should comply with the following channel class characteristics: ( 1 ) Sled or drop tower vehicle acceleration should be measured in accordance with the requirements of Channel Class 60, unless the acceleration is also integrated to obtain velocity or displacement, in which case, it should be measured in accordance with the Channel Class 180 requirements. ( 2 ) Belt restraint system loads s hould be measured in accordance with the requirements of Channel Class 60. ### Page 168 AC 27-1B, Chg 1 2/12/03 Page C - 60 ( 3 ) ATD head accelerations used for calculating the HIC should be measured in accordance with the requirements of Channel Class 1,000. ( 4 ) ATD femur forces may be measured if desired in accordance with Channel Class 600. ( 5 ) ATD pelvic/lumbar spinal colu mn force should be measured in accordance with the requirements of Channel Class 600. ( 6 ) The full-scale calibration range for each channel should provide sufficient dynamic range for the data being measured. ( 7 ) Digital conversion of analog data s hould provide sample resolution of not less than 1 percent of full-scale input. (B) Photographic instrumentation . Photographic instrumentation is used for documenting the response of the ATD and the test items to the dynamic test environment. Both high speed motion picture and still systems are used. ( 1 ) High-speed motion picture cameras that provide data used to calculate displacement or velocity should operate at a nominal speed of 1,000 pictures per second. Photo instrumentation methods should not be used for measurement of acceleration. The locations of the cameras and of targets or targeted measuring points within the field of view should be measured and documented. Targets should be at least 1/100 of the field width covered by the camera and should be of contrasting colors or should contrast with their background. The center of the target should be easily discernible. Rectilinearity of the image should be documented. If the image is not rectilinear, appropriate correction factors should be used in the data analysis process. A description of photographic calibration boards or scales within the camera field of view, the camera lens focal length, and the make and model of each camera and lens should be documented for each test. Appropriate digital or serial timing should be provided on the image media. A description of the timing signal, the offset of timing signal to the image, and the means of correlating the time of the image with the time of electronic data should be provided. A rigorous, verified analytical procedure should be used for data analysis. ( 2 ) Cameras operating at a nominal ra te of 200 pictures per second or greater can be used to document the response of ATD and test items if measurements are not required. For example, actions such as movement of the pelvic restraint system webbing (lap safety belt) off of the ATD pelvis or movement of upper torso restraint webbing off of the ATD’s shoulder can be observed by documentation cameras placed to obtain a “best view” of the anticipated event. These cameras should be provided with appropriate timing and a means of correlating the image with the time of electronic data. ### Page 169 Page C - 61 ( 3 ) Still image cameras can be used to document the pretest installation and the posttest response of the ATD’s and the test items. At least four pictures should be obtained from different positions around the test items in pretest and posttest conditions. Where an upper torso restraint system is installed, posttest pictures should be obtained before moving the ATD. For the posttest pictures, the ATD’s upper torso may be rotated to the approximate upright seated position so that the condition of the restraint system may be better documented, but no other change to the posttest response of the test item or ATD’s should be made. The pictures should document that the seat remained attached at all point of attachment to the test fixture. Still pictures can also be used to document posttest yielding of the seat for the purpose of showing that it would not impede the rapid evacuation of the airplane occupants. The ATD’s should be removed from the seat in preparation for still pictures used for that purpose. Targets or an appropriate target grid should be included in such pictures, and the views should be selected so that potential interference with the evacuation process can be determined. For tests where the ATD’s head impacts a fixture or another seat back, pictures should be taken to document the head contact areas. (iv) ATD . The tests discussed in this guidance were developed using modified forms of the ATD specified by the United States Code of Federal Regulations, Title 49, Part 572 Anthropomorphic Test Dummies, Subpart B – 50th Percentile Male. These “Part 572B” ATD’s were developed for automobile impact testing and have been shown to be reliable test devices capable of providing reproducible results in repeated testing. However, since ATD development is a continuing process, the standards allow use of “equivalent” dummies. See paragraph c.(2)(iv)(D) of this guidance. Dummy types should not be mixed when the tests discussed in this guidance are performed. (A) Modification for meas uring pelvic/lumbar column load. Since ATD’s have been developed for use in automobile testing to evaluate injury protection in forward, rearward, and sideward impacts, the ATD’s must be modified to measure the spinal load to comply with the § 27.562(c)(7). This load is influenced by a vertical direction component and by upper torso restraints which may produce a downward force component on the shoulders. To measure the load, a load (force) transducer is inserted into the ATD pelvis just below the lumbar column. This modification is shown in Figure 27.562-4. A commercially available “femur” load cell with end plates removed has been adapted to the modified ATD to measure the compression load between the pelvis and the lumbar spine column of the ATD. A “femur” load cell is commonly available to most test facilities and (according to specifications) is insensitive to bending and twisting moments. This feature prevents load transmission through the load cell as it measures the ATD lumbar/pelvis compression forces. To maintain the correct seated height of the ATD, the load cell is fixed in a rigid cup inserted into a hole bored in the top surface of the ATD pelvis, the top flange of which is bolted to the pelvis. If necessary, ballast should be added to the pelvis to maintain the specified weight of the assembly. Alternative approaches to measuring the axial force transmitted to the lumbar spinal column by the pelvis are acceptable if the method— ### Page 170 AC 27-1B, Chg 1 2/12/03 Page C - 62 ( 1 ) Accurately measures the axial forc e but is insensitive to moments and forces other than that being measured; ( 2 ) Maintains the intended alignment of the spinal column and the pelvis, the correct seated height, and the correct weight distribution of the ATD; and ( 3 ) Does not alter the other performan ce characteristics of the ATD. FIGURE AC 27.562-4 – Installation of Pelvic—Lumbar Spine Load Cell In Part 572B Anthropomorphic Dummy. (B) Figure 27.562-4 shows an acc eptable installation of a femur load cell (d) at the base of the ATD lumbar spine (a). The load cell is in line with the centerline of the lumbar spine and set below the top surface of the pelvis casting to maintain the seated height of the ATD. A rigid adapter cup (e) is fabricated to hold the load cell, and a hole is bored in the ATD pelvis to accept the cup. Provide clearance between the walls of the adapter cup and the load cell and the wires leading from the cell to avoid possible interference loads. The bottom of the load cell is bolted to the adapter cup. Adapter plates having similar hole patterns in their periphery are fabricated for the lower surface of the lumbar spine (b) and the upper surface of the load cell (c). These plates are fastened to the lumbar spine and load cell with screws through holes matching threaded holes in those components and are then joined together by bolts through the peripheral holes. The flange on the adapter cup has a bolt hole pattern matching that on the pelvis. The cup is fastened to the pelvis using screws to the threaded holes in the pelvis. Spacers (f) may be placed under the flange of the cup to obtain the specified ATD seating height. Additional weight should be placed in the cavity below the adapter ### Page 171 Page C - 63 cup to compensate for any weight lost because of this modification. The instrument cavity plug (g) is cut to provide clearance for the adapter cup and added weight. (C) Other ATD modifications . Flailing of the ATD arms often causes the “clavicle” used in the Part 572B ATD to break. To reduce the frequency of this failure, the clavicle may be replaced by a component having the same shape but made of higher strength material. This may increase the ATD weight slightly, but it would be acceptable for the tests discussed in this guidance. Another useful modification is the use of “submarining indicators” on the ATD pelvis. These electronic transducers are located on the anterior surface of the ilium of the ATD pelvis without altering its contour and indicate the position of the lap safety belt as it applies loads to the pelvis. Thus they can provide a direct record that the lap safety belt remains on the pelvis during the test and eliminates the need for careful review of high-speed camera images to make that determination. (D) Equivalent ATD . The continuing development of ATD for dynamic testing of seat restraint/crash-injury-protection systems is guided by goals of improved biofidelity (human-like response to the impact environment) and reproducibility of test results. The following criteria can be used to assess whether or not an ATD is equivalent to the present Part 572B ATD: ( 1 ) Fabrication in accordance with design and production specifications established and published by a regulatory agency responsible for crash injury protection systems; ( 2 ) Capability of providin g data for the measurements discussed in this guidance or of being readily altered to provide the data; ( 3 ) Evaluation by comparison with t he Part 572B ATD and shown to generate similar response to the impact environment discussed in this guidance; and ( 4 ) Any deviations from the Part 572B ATD configuration or performance are representative of the occupant of a civil aircraft in the impact environment discussed in this guidance. (E) Temperature and humidity . Since extremes of temperature and humidity can change the performance of ATD, the tests discussed in this guidance should be conducted at a temperature from 66° F to 78° F, and at a relative humidity from 10 percent to 70 percent. The ATD should have been maintained under these conditions for at least 4 hours prior to the test. (3) Test Preparation . Preparations for the tests should include selection of the test articles to be used in the tests, determination of the “most critical” conditions for the tests, and installation of the test articles, instrumentation, and ATD on the test fixture. Preparations pertaining to the normal operation of the test facility, such as safety ### Page 172 AC 27-1B, Chg 1 2/12/03 Page C - 64 provisions and the actual procedure for accomplishment of the tests, are particular to the test facility. These may be included in a test proposal or plan. (i) Selection of test articles . Many seat designs compose a “family or type” of seats which have the same basic structural design but differ in detail. For example, a basic seat frame configuration can allow for several different seat leg locations to permit installation in different airplanes. If these differences are of such a nature that their effect can be determined by rational analysis, then the analysis can determine the most highly stressed (“most critical”) configuration. The most highly stressed configuration would normally be selected for the dynamic tests so that the other configurations could be accepted by analysis and comparison with that configuration. The HIC depends on head impact (secondary impact after rotorcraft ground impact) and is more dependent on seat pitch for multiple row seats and on location for others than on seat structural stress for a given “family” of seats, so that the selection of the most highly stressed seat structure and the most critical seat pitch or location will permit these factors to be evaluated in one dual row test under the conditions of Test 2. Critical pelvic/lumbar spinal column forces are usually found under the vertical impact conditions of Test 1 but are influenced by the upper torso restraint in Test 2. Certain factors should be considered when employing that assumption. For example: (A) If the test item incorporat es some energy absorbing or load limiting design concept necessary to meet the test criteria or other requirement, a less severe loading condition may adversely affect the performance of that design concept as related to the pass-fail criteria. In such a case, it should be shown by rational analysis or additional testing that the design concept would continue to perform as intended even under the lower loads. (B) If different configuration of the same basic design incorporated load-carrying elements, especially joints or fasteners, which differed in detail design, the performance of each detail design should be demonstrated in a dynamic test. Experience has shown that small details in the design often cause problems in meeting the test performance criteria. (C) If structural strength is not the critical condition for achieving the performance criteria of the dynamic test, the true critical condition should be evaluated in a dynamic test. For example, if in one of the design configurations the restraint system attachment points are located so that the lap safety belt was more likely to slip above the ATD pelvis during the impact, then that configuration should also be dynamically tested even though the structural loading might be less. In all cases, the test item should be representative of the final production item in all structural elements and should include seat cushions, armrests and armcaps, functioning position adjustment mechanism, and correctly adjusted seat back breakover (if present), food trays or any other service or accoutrements required by the seat manufacturer or customer, and any other items of mass carried or positioned by the seat structure (e.g., weights simulating luggage carried or restrained by luggage restraint bars, fire ### Page 173 Page C - 65 extinguishers, survival equipment, etc.). If these items of mass are placed in a position that could limit the function of an energy absorbing design concept in the test item, they should be of representative shape and stiffness as well as weight. That is, seat stroking should perform properly when used in rotorcraft interiors. (ii) Considerat ion of test criteria. The test proposal or plan should be planned to achieve “most critical” conditions for the criteria that make up each test. (A) For multiple occupant seat a ssemblies, a rational structural analysis should be used to determine the number and seat location for the ATD and the direction for seat yaw in Test 2 to provide the most critical seat structural stress. This will usually result in unequally loaded seat legs. The seat deformation procedure should be selected to increase the load on the highest loaded seat leg and to stress the floor track or fitting in the most severe manner. The seat position in Test 2 depends on the upper torso restraint design. See c.(3)(ii)C below. (B) If multiple row testing is us ed to gather data for HIC in passenger seats, the seat pitch distance between seat rows should be selected within the allowable range, so that the head would be most likely to contact hard structure in the forward seat row. The effect of the 10° yaw in Test 2 and of any seat back breakover should be considered. Results from previous tests or rational analysis can be used to estimate the head strike path. Upper torso restraints may prevent head strike; however, leg kick loads into the front seat row require use of two rows. This kick load is a seat structural test not an ATD consideration. (C) If nonsymmetrical upper torso restraints (such as single diagonal shoulder belts) are used in a system, they should be installed on the test fixture in a position representative of that in the aircraft and that would most likely allow the ATD to move out of the restraint. For example, in a forward facing crew seat equipped with a single diagonal shoulder belt, the seat should be yawed in Test 2 in a direction such that the belt passes over the trailing shoulder. This is a part of the pass/fail criteria evaluation. (D) If a seat has sitting height adjustment, it should be tested in the highest position that could be used by a 50th percentile male occupant in the aircraft installation. See b.(3)(ii) of this guidance. (E) Floor deformation need not be considered in assessing the consequence of any seat deformation as related to the possible impairment of rapid evacuation of the airplane. After the test, the pitch and roll floor beams can be returned to their neutral position and the necessary measurements of the seat deformation made to determine the effect, if any, on rapid evacuation. (F) In some cases, it may not be possible to measure data for HIC during the test of the seat and torso restraint system. The design of the surrounding interior, such as the instrument panel, may not be known to the designer of the seat and torso ### Page 174 AC 27-1B, Chg 1 2/12/03 Page C - 66 restraint system, or the system may be used in several applications with different interior configurations. In such cases, it will be necessary to document the head strike path and the velocity along the path. This will require careful placement of photo instrumentation cameras and location of targets on the ATD representing the ATD head center of mass so that the necessary data can be obtained. These data can be used by the interior designer to ensure that head impact with the interior will not take place or that if possible head impact occurs, it will remain within the limits of the HIC. In the event the head impacts the specific interior, the interior under evaluation should be subjected to an individual special test to measure the head impact or HIC. The test is done using a rigid 6.5-inch diameter spherical head form weighing 15 pounds, (which includes necessary mass to represent the neck and a portion of the torso). The center of the head form is guided along the previously determined head strike path so that the form contacts the interior components at the velocity previously determined during the seat and torso restraint system dynamic test. Accelerometers located at the center of the head form would provide the data necessary for the HIC computation. If the interior component to be impacted by the ATD has significant inertial response to the impact environment, it will be necessary to evaluate those features or systems, such as breakover seatbacks or instrument panels designed to move forward, relative to the seat, in a dynamic test program which includes the full ATD occupant/seat/restraint system. See b.(3)(ii) of this guidance for ATD and panel location for adjustable crew seats. (iii) Use of ATD . ATD used in the tests discussed in this guidance should be maintained to perform in accordance with the requirements described in their specification. Periodic teardown and inspection of the ATD should be accomplished to identify and correct any worn or damaged components, and appropriate ATD calibration tests (as described in their specification) should be accomplished if major components are replaced. Each ATD should be clothed in form-fitting cotton stretch garments with short sleeves, mid-calf length pants, and shoes (size 11E) weighing about 2.5 pounds. The head and face of the ATD can be coated with chalk dust if it is desired to mark head contact areas on seats or other structure. The friction in limb joints should be set so that the joints barely restrain the weight of the limb when extended horizontally. The ATD should be placed in the seat in a uniform manner for reproducible test results. For the tests discussed in this guidance, the following procedures are adequate: (A) The ATD should be placed in the center of the seat in as nearly a symmetrical position as possible. (B) The ATD’s back should be agains t the seat back without clearance. This condition can be achieved if the ATD’s legs are lifted as it is lowered into the seat. Then, the ATD is pushed back into the seat back as it is lowered the last few inches into the seat pan. Once all lifting devices have been removed from the ATD, the ATD should be “rocked” slightly to settle it in the seat. (C) The ATD knees should be separated about 4 inches. ### Page 175 Page C - 67 (D) The ATD hands should be placed on the top of the legs, just behind the knees. If tests on crew seats are conducted in a mockup with aircraft controls, the ATD hands should be lightly tied to the controls. If only the seat and occupant restraint system are tested, the ATD hands should be tied together with a lack cord that provides about 24 inches of separation before the cord becomes tight. This will prevent excessive arm flail during the ATD rebound phase. (E) To the extent that they influence the injury criteria, all seat adjustments and controls should be in the design position intended for the 50th percentile male occupant. If seat and occupant restraint systems being tested are to be used in applications where requirements (placards) dictate particular positions for landing and takeoff, those positions should be used in the tests. (F) The feet should be in the appr opriate position for the type of seat tested (flat on the floor for a passenger seat or on control pedals or on a 45° footrest for flightcrew systems). The feet should be placed so that the centerlines of the lower legs are approximately parallel, unless the need for placing the feet on aircraft controls dictates otherwise. (iv) Installation of instrumentation . Professional practice should be followed when installing instrumentation. Care should be taken when installing the transducers to prevent deformation of the transducer body from causing errors in data. Lead-wires should be routed to avoid entanglement with the ATD or test item, and sufficient slack should be provided to allow motion of the ATD or test item without breaking the lead wires or disconnecting the transducer. Calibration procedures should consider the effect of long transducer lead-wires. Head accelerometer (transducer) should be installed in the ATD in accordance with the ATD specification and the instructions of the transducer manufacturer. The load cell between the pelvis and the lumbar spinal column should be installed as shown in Figure 27.562-4 of this guidance or in a manner that would provide equivalent data. (A) An upper torso restraint is required by § 27.785(b). The tension load should be measured in a segment of webbing between the ATD’s shoulders and the first contact of the webbing with hard structure (the anchorage point or a webbing guide). Restraint webbing should not be cut to insert a load cell in series with the webbing, since that would change the characteristics of the restraint system. Commercially available load cells can be placed over the webbing without cutting. They should be placed on free webbing and should not contact hard structure, seat upholstery, or the ATD during the test. They should not be used on double-reeved webbing, multiple-layered webbing, locally-stitched webbing, or folded webbing unless it can be demonstrated that these conditions do not cause errors in the data. These load cells should be calibrated using a length of webbing of the type used in the restraint system. If the placement of the load cell on the webbing causes the restraint system to sag, the weight of the load cell can be supported by light string or tape that will break away during the test. ### Page 176 AC 27-1B, Chg 1 2/12/03 Page C - 68 (B) Loads in restraint systems attachi ng directly to the test fixture can be measured by three-axis load cells fixed to the test fixture at the appropriate location. These commercially available load cells measure the forces in three orthogonal directions simultaneously, so that the direction as well as the magnitude of the force can be determined. If desired, similar load cells can be used to measure forces at other boundaries between the test fixture and the test item, such as the forces transmitted by the legs of the seat into the floor track. It is possible to use independent, single axis load cells arranged to provide similar data, but care should be taken to use load cells that can withstand significant cross-axis loading or bending without causing errors in the test data, or use careful (often complex) installation to protect the load cells from cross-axis loading or bending. Since load cells are sensitive to the inertial forces of their own internal mass and to the mass of fixtures located between them and the test article, as well as to forces applied by the test article, it may be necessary to compensate the test data for that inaccuracy if the error is significant. Data for such compensation will usually be obtained from an additional dynamic test replicating the load cell installation but will not include the test item. (v) Restraint system adjustment . The ATD should be sitting in the normal upright position. Care should be taken not to tighten the restraint system beyond the level reasonably expected in use and do not lock any emergency locking device (inertia reel) prior to the impact. Automatic locking retractors should be allowed to perform the webbing retraction and automatic locking function without assistance. Care should be taken that emergency locking retractors sensitive to acceleration do not lock prior to the impact test because of pre-impact acceleration applied by the test facility that is not present in a landing impact. If “comfort zone” retractors are used, they should be adjusted in accordance with instructions given to the user of the system. If manual adjustment of the restraint system is required, it should be sufficient to remove slack in the webbing, but it should not be adjusted so that it is unduly tight. Since the force required to adjust the length of the webbing can be as high as 11 pounds, a preload of 12-15 pounds is commonly recommended. This load is too small to be accurately measured by transducers selected to measure the high loads encountered in the impact test, so it should be measured manually as the restraint is being adjusted. Special gauges are commercially available to assist in this measurement. The preload should be checked and adjusted, if necessary, just prior to the floor deformation phase of the test. (vi) Repetition of tests . It may be necessary to repeat the tests discussed in this AC if accurate data are not collected in critical data channels or if some other error occurs (e.g., cameras fail to operate, impact pulse inadequate, etc.). Preparation for a repeated test should follow the same steps as for the initial test. The seat should be removed from the fixture, and its attachment fittings or floor track examined and replaced, if necessary, to correct any damage. The ATD should be carefully examined and repaired or adjusted, if necessary. It is usually preferable to use a new seat and restraint system for all repeated tests to preclude system failures due to undetected damage. A new seat and restraint system should be used if there is any detectable variation from the intended design configuration. ### Page 177 Page C - 69 d. Data Analysis And Compliance With The Criteria (1) General . All data obtained in the dynamic tests should be reviewed for errors. Baseline drift, “ringing,” and other common electronic instrumentation problems should be detected and corrected before the tests. Loss of data during the test is readily observed in a plot of the data vs. time and is typically indicated by sharp discontinuities in the data, often exceeding the amplitude limits of the data collection system. If these occur early in the test in essential data channels, the data should be rejected and the test repeated. If they occur late in the test, after the maximum data in each channel has been recorded, the validity of the data should be carefully evaluated, but the maximum values of the data may still be acceptable for the tests described in this guidance. The HIC does not represent a maximum data value, but represents an integration of data over a varying time base. The head acceleration measurements used for that computation should not be accepted if errors or loss of data are apparent in the data at any time from the beginning of the test until the ATD and all test articles are at rest after the test. (2) Impact pulse shape . Data for evaluating the impact pulse shape are obtained from an accelerometer that measures the acceleration in the direction parallel to the line of inertial response shown in Figure 27.562-1 of this guidance. The impact pulse intended for the tests discussed in this guidance has a symmetrical (isosceles) triangular shape. Since this ideal pulse is considered a minimum test condition, it is possible to evaluate the actual test pulse by comparing it with the ideal triangular pulse. The ideal pulse can be drawn to scale on the data plot of the test sled or carriage acceleration vs. time. The test pulse is acceptable if the plotted data are equal to or greater than the ideal impact pulse. This method can lead to a practical necessity of exceeding the ideal pulse by a significant degree, unless the test facility has precise control in generating the test pulse. A graphic technique may be used to evaluate test impact pulse shapes that are not precise isosceles triangles. A graphic technique is contained in paragraph f. (1) of this guidance. (3) Head Injury Criterion (HIC) . Data for determining the HIC need to be collected during the tests discussed in this guidance only if the ATD’s head is exposed to secondary impact. The HIC is a method for defining an acceptable limit; i.e., the maximum values of the HIC should not exceed 1,000 for head impact against broad interior surfaces in a crash. The HIC is reported as the maximum value, and the time interval during which the maximum value occurs is also given. Most facilities will make this computation if requested. The HIC is calculated by computer-based data analysis systems because manual attempts to use this method with real data are likely to be tedious. The HIC is calculated according to the following equation: ### Page 178 AC 27-1B, Chg 1 2/12/03 Page C - 70 Where: t1 and t2 are any two points in the time range during the head impact. The range should not exceed 0.050 seconds, and a(t) is the resultant head acceleration at the center of gravity (expressed in g’s) during the head form impact. (i) Data collection . The HIC is commonly based on data obtained from three mutually perpendicular accelerometers installed in the head of the ATD in accordance with the ATD specification. Data from these accelerometers are obtained using a data system conforming to Channel Class 1,000 as described in SAE Recommended Practice J211. For the tests discussed in this guidance (both ATD and head form), only the data taken during secondary head impact with the aircraft interior need be considered. Head impact is often indicated in the data by a rapid change in the magnitude of the acceleration. Alternately, a film of the test may show head impact which can be correlated with the acceleration data by using the time base common to both electronic and photographic instrumentation, or simple contact switches on the impacted surface can be used to define the initial contact time. (ii) HIC methodology . The following discussion outlines the basic method for computing the HIC. The magnitude of the resultant acceleration vector obtained from the three accelerometers is plotted against time. Then, beginning at the time of initial head contact (t1), the average value of the resultant acceleration is found for each increasing increment of time (t2 – t1), then integrating the curve between the range of t1 and t2 and then dividing the integral value by the time (t2 – t1). This calculation should use all data points provided by the minimum 8,000 samples per second digital sampling rate for the integration. However, the maximizing time intervals need be no more precise than 0.001 seconds. The average values are then raised to the 2.5 power and multiplied by the corresponding increment of time (t2 – t1). This procedure is then repeated, increasing t1 by 0.001 seconds for each repetition. The maximum value of the set of computations obtained from this procedure is the HIC. The procedure may be simplified by noting that the maximum value will only occur in intervals where the resultant magnitude of acceleration at t1 is equal to the resultant magnitude of acceleration at t2 and when the average resultant acceleration in that interval is equal to 5/3 times the acceleration at t1 or t2. (iii) Limitations . HIC does not consider injuries that can occur from contact with surfaces having small contact areas or sharp edges, especially if those surfaces are relatively rigid. These injuries can occur at low impact velocities, and are often described as “cosmetic” injuries; however, they can involve irreversible nerve damage and permanent disfigurement. While there is no generally accepted test procedure to provide quantitative assessment of these injuries, a judgmental evaluation of soft tissue injuries can be made by assessing tears or cuts in a synthetic skin placed ### Page 179 Page C - 71 over the ATD’s head or a head form during the test. Synthetic skins are discussed in the Society of Automotive Engineers Information Report SAE J202, Synthetic Skins for Automotive Testing. (4) Impact velocity . Impact velocity can be obtained by measurement of a time interval and a corresponding sled displacement occurring just before or after (for acceleration facilities) the test impact, and then dividing the displacement by the time interval. When making such a computation, the possible errors of the time and displacement measurements should be used to calculate a possible velocity measurement error, and the test impact velocity should exceed the velocity shown in Figure 27.562-1 by at least the velocity measurement error. If the sled is changing acceleration during the immediate pre-impact interval, or if the facility produces significant rebound of the sled, the effective impact velocity can be determined by integrating the plot of sled acceleration vs. time. If this method is used, the sled acceleration should be measured in accordance with Channel Class 180 requirements. (5) Upper torso restraint system load . The maximum load in the upper torso restraint system webbing can be obtained directly from a plot or listing of webbing load transducer output. If a three-axis load transducer, fixed to the test fixture, is used to obtain these data, the data from each axis should be combined to provide the resultant vector magnitude. If necessary, corrections should be made for the internal mass of the transducer and the fixture weight it supports. This correction will usually be necessary only when the inertial mass or fixture weight is high or when the correction becomes critical to demonstrate that the measurements fall below the specified limits. (6) Compressive load between the pelvis and lumbar column . The maximum compressive load between the pelvis and the lumbar column of the dummy can be obtained directly from a plot or listing of the output of the load transducer at that location. Since most load cells will indicate tension as well as compression, care should be taken that the polarity of the data has been correctly identified. (7) Retention of upper torso restraint straps. Retention of the upper torso restraint webbing straps on the ATD’s shoulders can be verified by observation of photometry or documentary camera coverage. The webbing should remain on the sloping portion of the ATD’s shoulder until the ATD rebounds after the test impact and the upper torso restraint straps are no longer carrying any load. The webbing straps should not bear on the neck or side of the head and should not slip to the upper rounded portion of the upper arm during that time period. (8) Retention of lap safety belt . Retention of the lap safety belt on the occupant’s (ATD) pelvis can be verified by observation of photometry or documentary camera coverage. The lap safety belt should remain on the ATD’s pelvis, bearing on or below each prominence representing the anterior superior iliac spines, until the ATD rebounds after the test impact and the lap safety belt becomes slack. If the lap safety belt does not become slack throughout the test, the belt should maintain the proper position throughout the test. Movement of the lap safety belt above the prominence is ### Page 180 AC 27-1B, Chg 1 2/12/03 Page C - 72 usually indicated by an abrupt displacement of the belt into the ATD’s soft abdominal insert which can be seen by careful observation of photo data from a camera located to provide a close view of the belt as it passes over the ATD’s pelvis. This movement of the belt is sometimes indicated in measurements of lap safety belt load (if such measurements are made) by a transient decrease or plateau in the belt force, as the belt slips over the prominence, followed by a gradual increase in belt force as the abdominal insert is loaded by the belt. Retention of the lap safety belt can also be verified by “submarining indicators” located on the ATD’s pelvis. These transducers are essentially a series of small, uncalibrated load cells placed in or above the rim of the ATD’s pelvis without changing its essential geometry. They indicate the position of the lap safety belt by producing an electrical signal when they are under load from the belt. (9) Femur load . Measuring femur loads is not required by the rotorcraft standards. If a seat is installed in an aircraft in a manner that will expose the system to loads from an occupant seated behind the seat system as well as the occupant seated in the seat system, the tests discussed in this guidance should be conducted in a manner to demonstrate that the system will perform properly under the combined loading. For example, Test 2 should be conducted with at least two rows of seats in place, as the seats in the first row carry the loads from the occupants in the first row, as well as the leg kick loads from the second row (also noted in c.(3)(ii)(A) of this guidance). (10) Seat attachment . Documentation that the seat and restraint system has remained attached at all points of attachment should be provided by still photographs that show the intact system components in the load path between the attachment points and the occupant. (11) Seat deformation . Occupant seats evaluated in the tests discussed in this guidance can deform permanently, either due to the action of discrete (impact) energy absorber systems included in the design or due to residual plastic deformation of their structural components. If this deformation is excessive, it could impede emergency evacuation. Each seat design may differ in this regard and should be evaluated according to its unique deformation characteristics. Permanent seat deformations are measured on the critically loaded seat subsequent to conduct of the tests required in § 27.562. The seat deformation is measured subsequent to completion of the dynamic tests and, where applicable, release of the applied pre-test floor deformation. (i) Seats . The following post-test deformations and limitations regarding emergency egress and access to exits may be used for showing compliance with § 27.785(j): (A) Forward or Rearward Directions . The forward or rearward deformations should not exceed a maximum of 4.0 inches (100 mm). In addition, the clearance between undeformed seat rows, measured as shown in Figure AC 27.562-5 (Dimension A), should be a minimum of 9.0 inches, except where seat rows lead to Type III or IV exits, where it should be a minimum of 11.0 inches. For seats with ### Page 181 Page C - 73 deformations exceeding 4.0 inches, the undeformed clearances between seats should be increased accordingly. In addition, at seat rows leading to Type III or IV exits, a minimum of 20 inches clearance, measured above the arm rests, must be maintained between adjacent seat rows. This measurement may be made with the seat backs returned, using no more than original seat back breakover forces, to their pretest upright or structurally deformed position. At other seat rows, the most forward surface of the seat back should not deform to a distance greater than one half of the original distance to the forwardmost hard structure on the seat (see Figure 27.562-6). (B) Downward Direction . There is no limitation on downward deformation, provided it can be demonstrated that the feet or legs of occupants seated aft would not be entrapped. Additionally, the seat bottom rotational deformation from the horizontal, measured at the centerline of each seat pan, should not exceed 20° forward (pitch down) or 35° aft (pitch up). This measurement should be made between the fore and aft extremities of the seat pan structure, considering the final position of the seat pan structure. In no case should rotation of the seat pan cause entrapment of the occupant. (C) Sideward Direction . ( 1 ) The deformed seat should not en croach more than 1.5 inches (40 mm) into the required space for longitudinal aisle at heights up to 25 inches (635 mm) above the floor. Determine which parts of the seat are at what heights prior to testing. ( 2 ) The deformed seat should not en croach more than 2.0 inches (50 mm) into the longitudinal aisle space at heights 25 inches (635 mm) or more above the floor. (D) Additional Considerations . In addition, none of the above deformations permit the seat to: ( 1 ) Affect the operation of any em ergency exit or encroach into an emergency exit opening for a distance from the exit not less than the width of the narrowest passenger seat installed. ### Page 182 AC 27-1B, Chg 1 2/12/03 Page C - 74 Measurement to be taken over full width of seat bottom cushion FIGURE AC 27.562-5 Post-test condition Pre-test condition Dimension “C” must be at least 50% of Dimension “B” FIGURE AC 27.562-6 ### Page 183 Page C - 75 ( 2 ) Encroach into any required passageway to large exits. ( 3 ) Encroach more than 1.5 inches into any cross aisle or evacuation (flight attendant) assist space for certain exits. (ii) Stowable Seats . Stowable seats, if used, should stow post-test and remain stowed without projecting into any required passageways. In addition, they should not project more than 1.5 inches into any flight attendant assist space or cross-aisle. (A) Seats that are Stowed Manually. A post-test stowage force no greater than 10 pounds (22kg) above the original stowage force may be used to stow the seat. (B) Seats that Stow Automatically . For a seat that may interfere with the opening of any exit, it shall automatically retract to a position that does not interfere with the exit opening as prescribed in § 27.807. For determining encroachment into passageways, cross-aisles, and assist spaces, a posttest stowage force no greater than 10 pounds (22kg), applied at a single point, may be used to assist automatic retraction. e. Test Documentation . (1) General . The tests discussed in this guidance should be documented in reports describing the test procedures and results. The test proposal, a description of the required tests, approved by the FAA should be referenced in the test report and contain the following: (i) Facility data . (A) The name and address of the test facility performing the tests. (B) The name and telephone number of the individual at the test facility responsible for conducting the tests. (C) A brief description and/ or photograph of each test fixture. (D) The date of the last instru mentation system calibration and the name and telephone number of the person responsible for instrumentation system calibration. (E) A statement confirming that the data collection was done in accordance with the recommendations in this guidance or a detailed description of the actual calibration procedure used and technical analysis showing equivalence to the recommendations of this AC ( Paragraph c.(2)(iii)(A)). ### Page 184 AC 27-1B, Chg 1 2/12/03 Page C - 76 (F) Manufacturer, governing specific ation, serial number, and test weight of ATD used in the tests, and a description of any modifications or repairs performed on the ATD which could cause them to deviate from the specification. (G) A description of the photogr aphic-instrumentation system used in the tests (Paragraph c.(2)(iii)(B)). (ii) Seat/Restraint system data . (A) Manufacturers name and identifying model numbers of the seat/restraint system used in the tests, with a brief description of the system, including identification and a functional description of all major components and photographs or drawings as applicable. (B) For unsymmetrical systems, an analysis supporting the selection of most critical conditions used in the tests. (2) Test Proposal or Plan and Description . The description of the test should be documented in enough detail so that the tests could be reproduced by following the guidance given in the report. The procedures outlined in this guidance can be referenced in the report but should be supplemented, as necessary, to describe the unique conditions of the individual seat design. (i) Pertinent dimensions and other details of the installation not included in the drawings of the test items should be provided. This can include footrests, restraint system webbing guides and restraint anchorages, “interior surface” simulations, bulkhead or sidewall attachments for seats or restraints, etc. (ii) The floor deformation procedur e, guided by goals of most critical loading for the test articles, should be documented. (iii) Placement and characteri stics of electronic and photographic instrumentation chosen for the test, beyond that information provided by the facility, should be documented. This can include special targets, grids or marking used for interpretation of photo documentation, and transducers and data channel characteristics for lap belt loads, floor reaction forces, or other measurements beyond those discussed in this guidance. (iv) Any unusual or unique activity or event pertinent to conducting the test should be documented. This could include use of special “breakaway” restraints or support for the ATD’s, test items or transducers, operational conditions or activities such as delayed or aborted test procedures, and failures of test fixtures, instrumentation system components or ATD. (3) Test results report . The documentation should include copies of all test results, analysis, and conclusions. As a minimum, the following should be documented: ### Page 185 Page C - 77 (i) Impact pulse shape (Paragraph d.(2)). (ii) HIC results for all ATD exposed to secondary head impact with interior components of the rotorcraft (Paragraph d.(3)), or head strike paths and velocities if secondary head impact is likely for future use in unique interiors (Paragraph c.(3)(iii)). (iii) Impact velocity (Paragraph d.(4)). (iv) Upper torso restraint system load if applicable (Paragraph d.(5)). (v) Compressive load between the pelvis and the lumbar column (Paragraph d.(6)). (vi) Retention of upper torso restraint straps if applicable (Paragraph d.(7)). (vii) Retention of lap safety belt (Paragraph d.(8)). (viii) Femur thigh loads, optional measurement. (ix) Seat attachment (Paragraph d.(10)). (x) Seat deformation (Paragraph d.(11)). (xi) Seat attachment reac tion time histories (Paragraph f.). (4) Dynamic Impact Test – Pass/Fail Criteria : The dynamic impact tests should demonstrate that: (i) The seat structure remains inta ct that is attached to the tracks or fittings, etc. (ii) The occupant retention syst em is capable of carrying the dynamic loads. (iii) The seat permanent deformati ons are within defined limits and will not significantly impede an occupant from releasing the torso restraints, standing and exiting the seat. (iv) If the ATD’s head is exposed to impact during the test, a HIC of 1,000 is not exceeded. Data may be obtained for use with other unique installations. (v) Where upper torso restraint straps are used, tension loads in individual straps do not exceed 7.78 kN (1,750 lbs.). If dual straps are used for ### Page 186 AC 27-1B, Chg 1 2/12/03 Page C - 78 restraining the upper torso, the total strap tension load does not exceed 8.90 kN (2,000 lbs.). (vi) The maximum compressive load measured between the pelvis and the lumbar column of the (ATD) does not exceed 6.67 kN (1,500 lbs.). (vii) Each upper torso restraint strap remains on the ATD shoulder during impact. (viii) The pelvic restraint rema ins on the ATD pelvis during impact. f. Procedures for Eval uating Impact Pulse Shapes. (1) Acceptable Evaluation Method . Data for evaluating the impact pulse shape are obtained from an accelerometer which measures the acceleration on the test fixture or sled at the seat location or equivalent location in the direction parallel to the line of inertial response shown in Figure 27.562-1 of this guidance. The impact pulses intended for the tests discussed in this guidance have an isosceles triangle shape. These ideal pulses are considered minimum test conditions. Since the actual acquired test pulses will normally differ from the ideal, it may be necessary to evaluate the acquired test pulses to insure the minimum requirements are satisfied. (2) An acceptable method to evaluat e the pulse shape should use the following steps: (i) Extend the calibra tion baseline (zero G) (ii) Locate the maximum peak deceleration (G p) indicated on the plot. (iii) Construct reference lines para llel to the baseline at levels of 0.1 Gp, 0.9 Gp, and 1.0 Gp. (iv) Construct an onset line thr ough the intersection points of the 0.1 Gp and 0.9 Gp reference lines with the increasing (onset) portion of the data plot. The data plot should not return to zero G between the two points selected. (v) Locate the intersection points of the onset line with the baseline and with the 1.0 Gp reference line. The interval between these two points, measured along the time axis of the data plot, is considered the rise time (tr) of the test impact pulse. (vi) The rise time of the test im pact pulse should not exceed the value of (tr) given in Figure AC 27.562-1 for each test. (vii) The area under the data plot cu rve within the rise time of the test impact pulse, Vra, should represent at least one half of the impact velocity given in Figure AC 27.562-1 for each test. If the value of peak acceleration measured in the test The text continues at https://openairdocs.org/ac/27-1B-CHG-1-2.md?from=187 --- 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.