# AC 27-1B CHG 1-9: Certification of Normal Category Rotorcraft - Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it) - Status: Current - Approved: June 23, 2023 - Last reviewed: June 29, 2023 - Family: AC 27-1 - Cancels: [AC 27-1B Change 8](https://openairdocs.org/ac/27-1B-Change-8) - 14 CFR parts: 27 - 14 CFR sections: [27.87](https://www.ecfr.gov/current/title-14/section-27.87), [27.903](https://www.ecfr.gov/current/title-14/section-27.903), [27.1305](https://www.ecfr.gov/current/title-14/section-27.1305), [27.1309](https://www.ecfr.gov/current/title-14/section-27.1309), [27.1329](https://www.ecfr.gov/current/title-14/section-27.1329), [27.1335](https://www.ecfr.gov/current/title-14/section-27.1335), [27.1353](https://www.ecfr.gov/current/title-14/section-27.1353), [27.1543](https://www.ecfr.gov/current/title-14/section-27.1543), [27.1545](https://www.ecfr.gov/current/title-14/section-27.1545), [27.1549](https://www.ecfr.gov/current/title-14/section-27.1549), [27.1555](https://www.ecfr.gov/current/title-14/section-27.1555), [27.1587](https://www.ecfr.gov/current/title-14/section-27.1587) - Pages: 1196 - PDF: https://openairdocs.org/files/sha256/5d/f4/5df495d139e348a84ea483664c33a83605903d77496646f9e1f1735d2c5bd584 - FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/DRSDOCID191065281520240304144337.0001 - DRS record last changed: November 21, 2025 - Page: https://openairdocs.org/ac/27-1 ## 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. ### 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 lllgblly slgral b)' **VICTOR W WICKLUND** o.1D: 2023.08.23 16;38;31 -04'00' 6/23/23 Advisory U.S. Department of Transportation Federal Aviation Circular Administration Subject: Certification of Normal Category Date: AC No: 27-1B Rotorcraft Initiated By: AIR-600 Change: 9 **1. PURPOSE.** a. This revision makes needed changes to this Advisory Circular (AC). b. The change number and the date of the changed material are shown at the top of each page. The vertical lines in the outside margin 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. PRINCIPAL CHANGES.** This Change 9 includes changes, edits, and additions for these AC sections: 27.87, 27.903, 27.1305, 27.1309, 27.1329, 27.1335, 27.1353, 27.1543, 27.1545, 27.1549, 27.1555, 27.1587, AC 27 Appendix B. The contents of this guidance document do not have the force and effect of law and are not meant to bind the public in any way. This guidance document is intended only to provide clarity to the public regarding existing requirements under the law or agency policies. **3. WEBSITE AVAILABILITY.** To access this AC electronically, you can go to either of the following sites: http://www.faa.gov/regulations_policies/advisory_circulars/ https://drs.faa.gov/browse Victor Wicklund, Acting Director, Policy & Standards Division Aircraft Certification Service ### Page 3 Advisory U.S. Department of Transportation Circular Federal Aviation Administration Subject: Certification of Normal Category Date: 06/29/2018 AC No: 27-1B Rotorcraft Initiated By: AIR-680 Change: 8 **1. PURPOSE.** a. This Advisory Circular (AC) publishes needed changes to the existing Advisory Circular (AC) material. b. The change number and the date of the changed material are shown at the top of each page. The vertical lines in the outside margin 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. PRINCIPAL CHANGES. This change 8 includes changes, edits, and additions for section 27.773. PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated D-63 thru D-66 9/30/1999 D-63 thru D-66 06/29/18 3. WEBSITE AVAILABILITY. To access this AC electronically, go to the Advisory Circulars library at http://www.faa.gov/regulations_policies/advisory_circulars/. Dr. Michael C. Romanowski Director, Policy and Innovation Division Aircraft Certification Service ### Page 4 0 Advisory U.S. Departmen1 of Transportation CircularFederal Aviation Ad,nlnistrotton Subject: CERTIFICATION OF NORMAL Date: FEB 4 2016 AC No: 27-1B CATEGORY ROTORCRAFT Initiated by: ASW-1 IO Cban~e: 7 **1. PURPOSE.** a. This change publishes needed changes to the existing material and adds Miscellaneous Guidance (MG) sections to the Advisory Circular (AC). b. The change number and the date of the changed material are shown at the top of each page. The vertical lines in the outside margin 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. PRINCIPAL CHANGES.** a. This change 7 includes changes, edits, and additions for these AC sections: 27.865B, 27.1321. 27 MG 1, 27 MG 16, and 27 Appendix B. b. This change adds to this AC ihe following MG sections: AC 27 MG 19 (Guidance on Electronic Display Systems (EDS) for Rotorcraft Installations), AC 27 MG 20 (Human Factors (HF)), and AC 27 MG 21 (Guidance on Creating a System Level Functional Hazard Assessment (FHA)). 3. WEBSITE AVAILABILITY. To access this AC electronically, go to the Advisory Aircraft Certification Service Circular 1library at http://www·.faa.gov/regulations policies/advisory circulars/. f~ a , Rotorcraft Directorate, ### Page 5 U.S. Deportment of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF NORMAL Date: 7/25/2014 AC No: 27-1B CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 6 **1. PURPOSE.** a. This Advisory Circular (AC) change publishes the new guidance on automatic flight guidance and control systems (AFGCS) installation in part 27 rotorcraft. b. 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. This change adds the miscellaneous guidance section 23 (MG 23), Automatic Flight Guidance and Control Systems (AFGCS) Installation in Part 27 Rotorcraft., to the AC. 3. WEB SITE AVAILABILITY. To access this AC electronically, go to the Advisory Circulars library at http://www.faa.gov/regulations_policies/advisory_circulars/. /s/ Lance T . Gant, for K im Smith Manager, Rotorcraft Directorate, Aircraft Certification Service ### Page 6 U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF NORMAL Date: 6/13/2014 AC No: 27-1B CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 5 **1. PURPOSE.** a. This Advisory Circular (AC) change publishes the new guidance on analyzing an advanced flight controls (AdFC) system installation in normal category rotorcraft. b. 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. This change adds the miscellaneous guidance section 17 (MG 17), Guidance on Analyzing an Advanced Flight Controls (AdFC) System, to the AC. 3. WEBSITE AVAILABILITY. To access this AC electronically, go to the Advisory Circulars library at http://www.faa.gov/regulations_policies/advisory_circulars/. /s/ Kim Smith Kim Smith Manager, Rotorcraft Directorate, Aircraft Certification Service ### Page 7 AC 27.1B, Chg 4 5/1/2014 Advisory Circular U.S Deportment of Transportation Federal AvkJtion Administration Subject: CERTIFICATION OF NORMAL Date: 5/1/2014 AC No: 27-1B CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 4 **1. PURPOSE.** a. This Advisory Circular (AC) publishes needed changes to the existing AC material. Additionally, there is incorporation of previously approved AC material and non-technical editorial changes to various sections. b. The change number and the date of the changed material are shown at the top of each page. The vertical lines in the outside margin indicate the beginning and end of each change. Pages that have different page numbers, but no text changes, will retain the previous heading information. 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. **2. PRINCIPAL CHANGES.** a. The AC sections previously approved on September 17, 2009 and posted separately to the Regulatory and Guidance Library (RGL), which are related to the Rotorcraft Performance and Handling Qualities rulemaking, are incorporated in this change 4. Those AC sections that were added or changed are: 27.25A, 27.49, 27.51A, 27.71, 27.73, 27.75A, 27.79, 27.87, 27.143A, 27.173A, 27.175A, 27.177A, 27.903B, 27.1587A, and 27 Appendix B. b. The changed AC 27 Miscellaneous Guidance (MG) 5 section (Agricultural Dispensing Equipment Installation), previously approved on December 15, 2009 and posted separately to the Regulatory and Guidance Library (RGL), is incorporated in this change 4. c. The AC 27.573 section previously approved on December 1, 2011 and posted separately to the Regulatory and Guidance Library (RGL), which is related to the Damage Tolerance and Fatigue Evaluation of Composite Rotorcraft Structures rulemaking, is incorporated in this change 4. Page ii ### Page 8 d. The changed AC 27 MG 6 section (Emergency Medical Service (EMS) Systems Installations Including: Interior Arrangements, Equipment, Helicopter Terrain Awareness and Warning System (HTAWS), Radio Altimeter, and Flight Data Monitoring System), previously approved on February 27, 2014, is incorporated in this change 4. e. The changed AC 27 MG 18 section (Helicopter Terrain Awareness and Warning System (HTAWS)), previously approved on February 27, 20014, is incorporated in this change 4. f. This change 4 also includes changes, edits, and additions for these AC sections: 27.29, 27.45, 27.75, 27.79, 27.141, 27.143, 27.151, 27.337, 27.351, 27.561, 27.562, 27.610, 27.783, 27.863, 27.865, 27.865B, 27.871, 27.903B, 27.939, 27.1011, 27.1093, 27.1093A, 27.1187, 27.1309, 27.1316, 27.1317, 27.1329, 27.1337, 27.1357, 27.1435. 27.1501, 27.1543, 27.1549, 27 MG 8, 27 MG 10, 27 MG 13, and 27 MG 22. 3. WEBSITE AVAILABILITY. To access this AC electronically, go to the Advisory Circulars library at http://www.faa.gov/regulations_policies/advisory_circulars/. /s/ Lance T. Gant, for Kim Smith Manager, Rotorcraft Directorate, Aircraft Certification Service Page iii ### Page 9 U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF NORMAL Date: 9/30/2008 AC No: 27-1B CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 3 **1. PURPOSE.** a. This Advisory Circular (AC) publishes needed changes to the existing AC material as a result of a safety-focused study. b. This change revises existing material in 6 sections. 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 or left margin indicates 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. Sections 27.571, 27.679, 27.695, 27.783, 27.901A, and 27.1351 are revised. 3. WEBSITE AVAILABILITY. To access this AC electronically, log on to http://www.airweb.faa.gov/rgl and then click on AC’s. ### Page 10 AC 27.1B, Chg 4 5/1/2014 AC 27-1B, Chg 3 PAGE CONTROL CHART Remove Pages Dated v thru xii ............................ 9/30/99 & 4/25/06 C-83 thru C-87 .................. 9/30/99 & 2/12/03 D-39 thru D-40 ...................................9/30/99 D-45 thru D-46 ...................................9/30/99 D-71 thru D-72 ...................................9/30/99 E-5 thru E-6 ....................................... 9/30/99 F-53 thru F-56................................ ....9/30/99 Signed by Scott A. Horn for Mark R. Schilling Acting Manager, Rotorcraft Directorate Aircraft Certification Service Insert Pages Dated v thru xii ............ 9/30/99, 4/25/06, & 9/30/08 C-83 thru C-88................................. 9/30/08 D-39 thru D-40.2.............................. 9/30/08 D-45 thru D-46................. 9/30/99 & 9/30/08 D-71 thru D-72.2.............. 9/30/99 & 9/30/08 E-5 thru E-6 ..................................... 9/30/08 F-53 thru F-56................................ .. 9/30/08 Page ii ### Page 11 U.S. Department of Transportation Federal Aviation Administration Advisory Circular ________________________________________ ______ Subject: CERTIFICATION OF NORMAL Date: 4/25/06 AC No: 27-1B CATEGORY ROTORCRAFT Initiated ASW-110 Change 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 13 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 MG 12 has been revised by a working group and is now contained in AC 27.865B, in Subpart D. c. New paragraph MG 18, Helicopter Terrain and Awareness Systems (HTAWS), is added to Chapter 3. d. New figure AC 27.351-2 is added in Chapter 2. e.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 12 I I I AC 27.1B, Chg 4 5/1/2014 AC 27-1B, Chg 2 PAGE CONTROL CHART Remove pages Dated v thru x .............................. 9/30/99 & 2/12/03 xiii thru xxi......................................... 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-109 thru D-112.............. 9/30/99 & 2/12/03 F-33 thru F-46.................................... 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 David A. Downey Manager, Rotorcraft Directorate Aircraft Certification Service Insert Pages Dated v thru x................................................4/12/06 xiii thru xxii ........................................4/12/06 C-9 thru C-87......................................4/12/06 D-1 thru D-58 .....................................4/12/06 D-67 thru D-70 ...................................4/12/06 D-109 thru D-133 ...............................4/12/06 F-33 thru F-46.....................................4/12/06 G-43 thru G-49 ...................................4/12/06 MG 1-15 thru MG 1-16 ......................4/12/06 MG 4-9 thru MG 4-13 ........................4/12/06 MG 8-1 thru MG 8-31 ........................4/12/06 MG 12-1 .............................................4/12/06 MG 16-1 thru MG 16-18 ....................4/12/06 MG 17-1 .............................................4/12/06 MG 18-1 thru MG 18-27 ....................4/12/06 Apdx A-47 thru Apdx A-71................4/12/06 Apdx B-11 thru Apdx B-20 ................4/12/06 Page ii ### Page 13 U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF NORMAL Date: 2/12/03 AC No: 27-1B CATEGORY ROTORCRAFT Initiated ASW-110 Change 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. f. Appendix A (previously reserved) is added. g. 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. h. 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. ### Page 14 AC 27.1B, Chg 4 5/1/2014 4. WEBSITE AVAILABILITY. To access this AC electronically, log on to http://www.airweb.faa.gov/rgl and then click on AC’s. AC 27-1B, Chg 1 PAGE CONTROL CHART Remove 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 David A. Downey Manager, Rotorcraft Directorate Aircraft Certification Service Insert Pages Dated 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 Page ii ### Page 15 0 Advisory Circular ______________________________________________ U.S. Department of Transportation Federal Aviation Administration Subject: CERTIFICATION OF NORMAL Date: 9/30/99 AC No: 27-1B CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 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. ### Page 16 AC 27.1B, Chg 4 5/1/2014 • 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. 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. Page ii ### Page 17 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 iii ### Page 18 CONTENTS Page No. **CHAPTER 1 - PART 21 CERTIFICATION PROCEDURES FOR** PRODUCTS AND PARTS 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 and Correction Report .............FAR 21-6 CHAPTER 2 - PART 27 AIRWORTHINESS STANDARDS: **NORMAL CATEGORY ROTORCRAFT** **SUBPART A** **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 27.25A Weight Limits … B-4 AC 27.27 Center of Gravity Limits..................................................................... B-4.2 AC 27.29 Empty Weight and Corresponding Center of Gravity ........................ B-6.2 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 27.45C Performance – General … B-18 Page No. Page iii ### Page 19 AC 27.1B, Chg 9 6/23/23 AC 27.49 Performance at Mini mum Operating Speed ...................................... B-24.2 AC 27.51 Takeoff … B-25 AC 27.51A Takeoff … B-26 AC 27.65 AC 27.65A AC 27.67 AC 27.71 AC 27.73 AC 27.75 AC 27.75A AC 27.79 AC 27.79A AC 27.87 AC 27.87A AC 27.141 AC 27.143 AC 27.143A AC 27.151 AC 27.161 AC 27.171 AC 27.173 AC 27.173A AC 27.175 AC 27.175A AC 27.177 AC 27.177A AC 27.231 AC 27.235 AC 27.239 AC 27.241 AC 27.251 Page iv SUBPART B -FLIGHT (continued) PERFORMANCE (continued) Climb: All Eng ines Operating ........................................................... B-26.2 Climb: All Engines Operating … B-28 Climb: One Engine Inoperative … B-28 Autorotation Performance … B-29 Performance at Minimum Operating Speed … B-32 Landing ............................................................................................. B-34.1 Landing … B-35 Limiting Height - Speed Envelope … B-37 Limiting Height - Speed Envelope … B-45 Height-Speed Envelope … B-45 Height-Velocity Envelope .................................................................. B-45.1 **FLIGHT CHARACTERISTICS** Flight Characteristics - General … B-46 Controllability and Maneuverability … B-48 Controllability and Maneuverability.................................................... B-52.2 Flight Controls................................................................................... B-52.8 Trim Control … B-53 Stability: General … B-53 Static Longitudinal Stability … B-54 Static Longitudinal Stability … B-55 Demonstration of Static Longitudinal Stability … B-58 Demonstration of Static Longitudinal Stability ................................... B-60.2 Static Directional Stability.................................................................. B-60.6 Static Directional Stability.................................................................. B-60.7 **GROUND AND WATER HANDLING CHARACTERISTICS** Ground and Water Handling Characteristics - General … B-61 Taxiing Condition … B-61 Spray Characteristics … B-62 Ground Resonance … B-63 **MISCELLANEOUS FLIGHT REQUIREMENTS** Vibration … B-68 ### Page 20 Page No. **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 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 Page v ### Page 21 AC 27.1B, Chg 4 5/1/2014 Page No. SUBPART C - STRENGTH REQUIREMENTS (continued) GROUND LOADS (continued) 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 AC 27.497 Ground Loading Conditions: Landing Gear With Tail Wheels … C-32 AC 27.501 Ground Loading Conditions: Landing Gear With Skids … C-34 AC 27.501A Ground Loading Conditions: Landing Gear With Skids … C-35 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-44 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 AC 27.573 Damage Tolerance and Fatigue Evaluation of Composite … C-88 Rotorcraft Structures Page vi ### Page 22 Page No. **SUBPART D - DESIGN AND CONSTRUCTION** **GENERAL** 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-6 AC 27.609 Protection of Structure … D-6 AC 27.610 Lightning Protection … D-8 AC 27.611 Inspection Provisions … D-11 AC 27.613 Material Strength Properties and Design Values … D-13 AC 27.613A Material Strength Properties and Design Values … D-14 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, and Power-Operated Systems... D-36 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-40 AC 27.681 Limit Load Static Tests...................................................................... D-40.1 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 vii ### Page 23 AC 27.1B, Chg 4 5/1/2014 Page No. SUBPART D - DESIGN AND CONSTRUCTION (continued) **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-72.2 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 Page viii ### Page 24 AC 27-1B, Chg 9 AC 27.853 AC 27.853A AC 27.855 AC 27.859 AC 27.861 AC 27.861A AC 27.863 AC 27.865 AC 27.865A AC 27.865B AC 27.871 AC 27.873 AC 27.901 AC 27.901A AC 27.903 AC 27.903A AC 27.903B AC 27.903C AC 27.907 AC 27.917 AC 27.921 AC 27.923 AC 27.923A AC 27.923B AC 27.927 AC 27.927A AC 27.931 Page No. SUBPART D - DESIGN AND CONSTRUCTION (continued) **FIRE PROTECTION** Compartment Interiors … D-98 Compartment Interiors … D-99 Cargo and Baggage Compartments … D-99 Heating Systems … D-101 Fire Protection of Structure, Controls, and Other Parts … D-105 Fire Protection of Structure, Controls, and Other Parts … D-106 Flammable Fluid Fire Protection … D-107 **EXTERNAL LOADS** External Load Attaching Means … D-109 External Load Attaching Means … D-110 External Loads (previously MG 12) … D-111 **MISCELLANEOUS** Leveling Marks … D-133 Ballast Provisions … D-133 **SUBPART E - POWERPLANT** **GENERAL** Installation … E-1 Installation … E-5 Engines ............................................................................................. E-5.2 Engines … E-7 Engines … E-9 Engines ............................................................................................. E-10.1 Engine Vibration................................................................................ E-10.1a ROTOR DRIVE SYSTEM Design............................................................................................... E-10.2 Rotor Brake … E-11 Rotor Drive System and Control Mechanism Tests … E-12 Rotor Drive System and Control Mechanism Tests … E-17 Rotor Drive System and Control Mechanism Tests … E-18 Additional Tests … E-19 Additional Tests … E-21 Shafting Critical Speeds … E-21 Page ix 6/23/23 ### Page 25 AC 27-1B, Chg 4 5/1/2014 Page No. SUBPART E – POWERPLANT (continued) ROTOR DRIVE SYSTEM (continued) AC 27.935 Shafting Joints … E-26 AC 27.939 Turbine Engine Operating Characteristics … E-26 **FUEL SYSTEM** 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 A C 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 Page x ### Page 26 Page No. SUBPART E - POWERPLANT (continued) **OIL SYSTEM** AC 27.1011 Engines - 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 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-101 AC 27.1093A Induction System Icing Protection..................................................... E-112.1 **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 Page No. Page xi ### Page 27 C 27-1B, Chg 9 A 6/23/23 AC 27-1B Chg 96/23/23 S UBPART E - POWERPLANT (continued) POWERPLANT FIRE PROTECTION Lines, Fittings, and ComponentsAC 27.1183 AC 27.1185 AC 27.1187 AC 27.1189 AC 27.1191 AC 27.1193 AC 27.1193A AC 27.1194 AC 27.1195 AC 27.1301 AC 27.1303 AC 27.1305 AC 27.1305A AC 27.1305B AC 27.1305C AC 27.1307 AC 27.1309 AC 27.1309A AC 27.1309B AC 27.1316 AC 27.1317 AC 27.1321 AC 27.1322 AC 27.1323 AC 27.1325 AC 27.1327 AC 27.1329 AC 27.1329A AC 27.1329B AC 27.1335 AC 27.1335A AC 27.1337 AC 27.1337A . … E-124 Flammable Fluids … E-125 Ventilation and Drainage … E-126 Shutoff Means … E-127 Firewalls … E-128 Cowling and Engine Compartment Covering … E-132 Cowling and Engine Compartment Covering … E-133 Other Surfaces … E-134 Fire Detector Systems … E-135 **SUBPART F - EQUIPMENT** **GENERAL** Function and Installation … F-1 Flight and Navigation Instruments … F-1 Powerplant Instruments … F-7 Powerplant Instruments … F-8 Powerplant Instruments … F-9 Powerplant Instruments .................................................................... F-11.1 Miscellaneous Equipment … F-12 Equipment, Systems, and Installations … F-13 Equipment, Systems, and Installations … F-30 Equipment, Systems, and Installations ............................................ F-30.1 Electrical and Electronic System Lightning Protection … F-31 High-Intensity Radiated Fields (HIRF) Protection … F-32 INSTRUMENTS: INSTALLATION Arrangement and Visibility … F-34 Flight Director Systems ..................................................................... F-54.1 Powerplant Instruments .................................................................... F-54.2 Powerplant Instruments .................................................................... F-54.3 Warning, Caution, and Advisory Lights ............................................. F-34.2 Airspeed Indicating System … F-39 Static Pressure Systems … F-40 Magnetic Direction Indicator … F-41 Automatic Pilot System … F-41 Automatic Pilot System … F-51 Automatic Pilot System and Flight Guidance System. ...................... F-52.1 Flight Director Systems … F-53 Page xii ### Page 28 P age No. SUBPART F – EQUIPMENT (continued) ELECTRICAL SYSTEMS AND EQUIPMENT Electrical Systems and Equipment - GeneralAC 27.1351 AC 27.1353 AC 27.1353A AC 27.1357 AC 27.1361 AC 27.1365 AC 27.1367 AC 27.1381 AC 27.1383 AC 27.1385 AC 27.1387 AC 27.1389 AC 27.1391 AC 27.1393 AC 27.1395 AC 27.1397 AC 27.1399 AC 27.1401 AC 27.1411 AC 27.1413 AC 27.1415 AC 27.1419 AC 27.1435 AC 27.1457 AC 27.1459 AC 27.1461 AC 27.1501 . … F-55 Storage Battery Design and Installation … F-58 Energy Storage Systems ................................................................... F-62.1 Circuit Protective Devices ................................................................. F-62.13 Master Switch … F-64 Electric Cables … F-65 Switches … F-66 **LIGHTS** Instrument Lights … F-67 Rear Position Lights Rear Position Lights Rear Position Lights Landing Lights … F-67 Position Light System Installation … F-68 Position Light System Dihedral Angles … F-68 Position Light Distribution and Intensities … F-68 Minimum Intensities in the Horizontal Plane of Forward and … F-68 Minimum Intensities in Any Vertical Plane of Forward and … F-68 Maximum Intensities In Overlapping Beams of Forward and … F-68 Color Specifications … F-68 Riding Light … F-68 Anti-collision Light System … F-70 **SAFETY EQUIPMENT** General … F-71 Safety Belts … F-73 Ditching Equipment … F-74 Ice Protection … F-75 Hydraulic Systems … F-97 Cockpit Voice Recorder … F-99 Flight Recorders … F-101 Equipment Containing High Energy Rotors … F-103 **SUBPART G - OPERATING LIMITATIONS AND INFORMATION** General … G-1 Page xiii ### Page 29 AC 27.1B, Chg 9 AC 27.1503 AC 27.1505 A C 27.1509 AC 27.1519 AC 27.1521 AC 27.1521A AC 27.1521B AC 27.1523 AC 27.1525 AC 27.1527 AC 27.1529 AC 27.1529A AC 27.1541 AC 27.1543 AC 27.1543A AC 27.1545 AC 27.1545A AC 27.1547 AC 27.1549 AC 27.1549A AC 27.1549B AC 27.1549C AC 27.1551 AC 27.1553 AC 27.1555 AC 27.1555A AC 27.1557 AC 27.1559 AC 27.1561 AC 27.1565 Page No. SUBPART G - OPERATING LIMITATIONS AND INFORMATION (continued) **OPERATING LIMITATIONS** Airspeed Limitations: General … G-1 Never-Exceed Speed … G-4 Rotor Speed … G-6 Weight and Center of Gravity … G-8 Powerplant Limitations … G-9 Powerplant Limitations … G-10 Powerplant Limitations … G-10 Minimum Flightcrew … G-11 Kinds of Operations … G-15 Maximum Operating Altitude … G-16 Instructions for Continued Airworthiness (Maintenance Manual) … G-16 Instructions for Continued Airworthiness (Maintenance Manual) … G-20 **MARKINGS AND PLACARDS** Marking and Placards - General … G-21 Instrument Markings: General … G-21 Instrument Markings: General......................................................... G-25.1 Airspeed Indicator ............................................................................. G-25.6 Airspeed Indicator ............................................................................. G-25.7 Magnetic Direction Indicator.............................................................. G-25.9 Powerplant Markings......................................................................... G-25.10 Powerplant Instruments .................................................................... G-25.10 Powerplant Instruments .................................................................... G-25.10 Powerplant Instruments … G-26 Oil Quantity Indicators....................................................................... G-2 6.1 Fuel Quantity Indicator ...................................................................... G-26.2 Control Markings … G-27 Control Markings ............................................................................... G-28.1 Miscellaneous Markings and Placards .............................................. G-28.2 Limitations Placard … G-30 Saf ety Equipment … G-31 Tail Rotor … G-32 ROTORCRAFT FLIGHT MANUAL AND APPROVED MANUAL MATERIAL AC 27.1581 AC 27.1583 AC 27.1585 AC 27.1585A AC 27.1587 AC 27.1587A AC 27.1587B AC 27.1589 Rotorcraft Flight Manual - General … G-33 Operating Limitations … G-40 Operating Procedures … G-42 Operating Procedures … G-45 Performance Information … G-45 Performance Information … G-47 Performance Information … G-48 Loading Information ......................................................................... G-48.1 Page xiv 6/23/23 ### Page 30 Page No. **CHAPTER 3 - MISCELLANEOUS GUIDANCE (MG),** NORMAL CATEGORY ROTORCRAFT AC 27 MG 1 Certification Procedure for Rotorcraft Avionics Equipment............. MG 1-1 System Tests §§ 27.79(b)(2), 27.141(b)(1), and 27.143(d)(1) Including: Interior Arrangements, Equipment, Helicopter Terrain Awareness and Warning System (HTAWS), Radio Altimeter, and Flight Data Monitoring System. System Air Conditioning Systems Systems (HUMS) Systems (NVIS) Equipment System Installations Assessment (FHA) Installation in Part 27 Rotorcraft AC 27 MG 2 Standardized Test Procedure for Rotorcraft DC Electrical ............. MG 2-1 AC 27 MG 3 Definition of Engine Isolation Features as Applied to .................... MG 3-1 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 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 Flotation ......... MG 10-1 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 .................. MG 14-1 AC 27 MG 15 Airworthiness Approval of Rotorcraft Health Usage Monitoring ..... MG 15-1 AC 27 MG 16 Certification Procedure for Rotorcraft Night Vision Imaging .......... MG 16-1 AC 27 MG 17 Guidance on Analyzing an Advanced Flight Controls (AdFC) ........ MG 17-1 AC 27 MG 18 Helicopter Terrain Awareness and Warning System (HTAWS)...... MG 18-1 AC 27 MG 19 Guidance on Electronic Display Systems (EDS) for Rotorcraft ...... MG 19-1 AC 27 MG 20 Human Factors (HF)....................................................................... MG 20-1 AC 27 MG 21 Guidance on Creating a System Level Functional Hazard ............ MG 21-1 AC 27 MG 22 Rotorcraft One Engine Inoperative (OEI) Training Mode................ MG 22-1 AC 27 MG 23 Automatic Flight Guidance and Control systems (AFGCS ) ........... MG 23-1 Page xv ### Page 31 6/23/23AC 27-1B, Chg 9 APPE NDICES 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 xvi ### Page 32 FIGURES The first digits in the figure number are the same as the related section number. Figure No. Title Page AC 27.45-1 H-V Requirements … B-20 Temperature - Sea Level Standard Day 400 RPM Attached at Floor Level Part 572B Anthropomorphic Dummy Seat Bottom Cushion AC 27.562-6 Post Test Condition AC 27.45-2 IGE Controllability Requirements … B-21 AC 27.45-3 Shaft Horsepower vs. Turbine Outlet … B-22 AC 27.45-4 Uninstalled Takeoff Power Available … B-23 AC 27.45-5 Installed Takeoff Power Available, Anti-Ice Off, … B-24 AC 27.45-6 Power Assurance Check Chart .................................................... B-24.1 AC 27.71-1 Autorotational Characteristics - Typical … B-31 AC 27.79-1 Height-Velocity Diagram … B-43 AC 27.79-2 Altitude/Weight Accountability … B-44 AC 27.175-1 Static Longitudinal Stability .......................................................... B-60.1 AC 27.175A-1 Static Longitudinal Stability .......................................................... B-60.5 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 Deformation Fixture; Seat Legs … C-54 AC 27.562-3 Schematic Test Fixture; Side Wall Mounted Seat … C-58 AC 27.562-4 Installation of Pelvic-Lumbar Spine Load Cell in … C-62 AC 27.562-5 Measurement to be Taken Over Full Width of … C-74 C-74 Page xvii ### Page 33 AC 27.1B, Chg 4 5/1/2014 FIGURES (continued) Figure No. Title Page AC 27.562-7 Impulse Shape … C-80 Size Relationships for Fail-Safe Substantiation … C113 Actuator Support-Landing Gear Model XYZ Organization.............................................................................. Versus Applicable Cargo Type Data and Definition Summary Hose-End Fitting Combinations Calculations Load Calculations, Frangible Bolts Load Calculations, Frangible Baffle AC 27.562-8 Change in Velocity … C-80 AC 27.571-1 Comparison Percent of Time in Level Flight … C-85 AC 27.573-1. Schematic diagram of building block tests … C102 AC 27.573-2. Characterization of Impact Damage … C-107 AC 27.573-3. Residual Strength vs. Time … C-112 AC 27.573-4. Illustration of Residual Strength and Damage AC 27.621-1 Example of Casting Load Sheet Retract … D-24 AC 27.621-2 Example of Castings to be Static Tested for … D-26 AC 27.773-1 Cockpit Visibility … D-67 AC 27.801-1 Sea State Code - World Meteorological … D-90 AC 27.865B-1 U.S. Operational (Part 133) Rotorcraft-Load Combination … D-116 AC 27.931-1 Cantilevered Shaft, First Critical Speed … E-25 AC 27.931-2 Shaft Between Support Bearings, First Critical Speed … E-25 AC 27.952-1 Static Tension 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 and … E-52 AC 27.952-4 Typical Method of Breakaway Fuel Fitting Load … E-53 AC 27.952-5 Typical Methods of Frangible or Deformable Attachment … E-54 AC 27.952-6 Typical Methods of Frangible or Deformable Attachment … E-55 AC 27.955-1 Fuel Flow E-61 … E-61 Page xviii ### Page 34 FIGURES (continued) Figure No. Title Page 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 Failure Condition Effects AC 27.1309-1 Relationship between Probability and Severity of .........................F-30.5 AC 27.1309-2 Failure Condition Categories and Probably Definitions ...............F-30.6 Diameter Diameter Dispensing Equipment Design Valve for Fail-Safe Substantiation for Three Rotorcraft Types AC 27.1309-4 Safety Assessment Process Overview ........................................F-30.10 AC 27.1309-5 Depth of Analysis Flowchart ........................................................F-30.11 AC 27.1321-1 Primary Fields of View ................................................................. F - 34 AC 27,1321-2 Secondary Fields of View............................................................. F – 34.1 AC 27.1329-1 Height Loss Method … F-45 AC 27.1329-2 Deviation Profile … F-49 AC 27.1329-3 Deviation Profile … F-50 AC 27.1419-1 Continuous Icing-Temperature vs. Altitude Limits … F-93 AC 27.1419-2 Intermittent Icing-Temperature vs. Altitude Limits … F-94 AC 27.1419-3 Continuous Icing-Liquid Water Content vs Drop … F-95 AC 27.1419-4 Intermittent Icing-Liquid Water Content vs. Drop … F-96 AC 27 MG 3-1 Part 29 Engine Isolation Rules..................................................... MG 3-5 AC 27 MG 5-1 Acceptable Ultimate Load Factor for Agricultural ......................... MG 5-3 AC 27 MG 5-2 Sketch of Tank Pressure Test...................................................... MG 5-7 AC 27 MG 6-1 Typical Liquid Oxygen System..................................................... MG 6-21 AC 27 MG 6-2 Typical Liquid Oxygen System - Using Combination....................MG 6-22 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 AC 27 MG 8-3 Residual Strength vs. Time .......................................................... MG 8-23 AC 27- MG 8-4 Residual Strength and Damage Size Relationships.....................MG 8-24 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 Flight......................MG 11-7 AC 27 MG 11-2 Flight Regime to be Investigated Power-On Operation................MG 11-17 Page xix ### Page 35 AC 27.1B, Chg 7 xx/xx/2014 FIGURES (continued) Figure No. Title Page AC 27 MG 11-3 Flight Regime to be Investigated for Power-Off ...........................MG 11-17 Operation Investigated Reduction of Software Level Software with a Ground Base that uses COTS Software for Operational Software AC 27 MG 11-4 Simplified Analytical Method for Safe Life Evaluation ..................MG 11-18 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 Measurement Program Conditions to be...................MG 11-20 AC 27 MG 11-8 Lifting Helicopter 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 Applied for ....................MG 15-8 AC 27 MG 15-2 Flow Chart for Application of HUMS Specific ............................... MG 15-11 AC 27 MG 17-1 Verification Methodology.............................................................. MG 17-4 AC 27 MG 19-1 Primary Field of View ................................................................... MG 19-15 AC 27 MG 19-2 Secondary Field of View .............................................................. MG 19-16 AC 27 MG 20-1 Primary and Secondary Horizontal Fields of View .......................MG 20-18 AC 27 MG 20-2 Primary and Secondary Vertical Fields of View............................MG 20-19 AC 27 MG 20-3 Sample HF Compliance Matrix..................................................... MG 20-27 AC 27 MG 21-1 Generic Examples by ATA ........................................................... MG 21-5 AC 27 APX B-1 Rotorcraft Dynamic Stability Requirements for IFR......................Apdx B-27 Page xx ### Page 36 Acronyms AC ACO ADF ADI AEH AEO AdFC AFCS AFGCS AGL AHRS Amdt. APU ATC BIM CAM CAR CAS CBIM CDP CFR CG COTS CPS CRFS CRI CRT CVR DER DME EASA ECAS ECU EDS advisory circular aircraft certification office automatic direction finding attitude direction indicator airborne electronic hardware all engines operating advanced flight control automatic flight control systems automatic flight guidance and control systems above ground level attitude heading reference system Amendment auxiliary power unit air traffic control blade inspection method cockpit area microphone Civil Air Regulations calibrated air speed cockpit blade inspection method critical decision point Code of Federal Regulations center of gravity commercial off-the-shelf cycles per second crash resistant fuel system certification review item cathode ray tube cockpit voice recorder designated engineering representative distance measuring equipment European Aviation Safety Agency engine caution advisory systems environmental control unit electronic data system EFIS electronic flight instrument System[Replaced by EDS.] EFP engine failure point EHE exhaust heat exchanger ELOS equivalent level of safety ELT emergency locator transmitter EMC electromagnetic compatibility EMI electromagnetic interference EMS emergency medical service EOL end of life EUROCAE European Organization for Civil Aviation Equipment EVS enhanced vision system FAA Federal Aviation Administration FADEC full authority digital engine control FAR Federal Aviation Regulations [OBSOLETE term.] FEM finite element modeling FHA functional hazard assessment FMEA failure mode and effects analysis FPM feet per minute FPS feet per second FTA fault tree analysis GPS global positioning system GW gross weight HEC human external cargo 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 Page xxi ### Page 37 AC 27.1B, Chg 4 5/1/2014 PCDS person nel carrying device IGE in ground effect Systems 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 LPV localizer performance with vertical guidance LRU line replaceable unit LWC liquid water content MC maximum continuous MCP maximum continuous power MEL minimum equipment list MG miscellaneous guidance MGB main gearbox MGT measured gas temperature MGW maximum gross weight MMEL master minimum equipment list MSL mean sea level MVD median volume diameter NASA National Aeronautics and Space Administration NDI non-destructive inspection NM nautical mile NPRM notice of proposed rulemaking NTSB National Transportation Safety Board NVG night vision goggles NVIS night vision imaging systems OAT outside air temperature OEI one engine inoperative OGE out of ground effect PBA pitch bias actuator Page xxii PCF PIO PSA PSIG QPL RFM RFMS RGL RPM RTCA RVR SAE SAR SAS SCAS SHP S/N SRM SSA STC STOL TBO TC TCDS TDP TIA TIR TOP TOT TSO TVP VBIM VFR VMC post crash fire pilot induced oscillation preliminary safety assessment pounds per square inch gauge qualified parts list rotorcraft flight manual rotorcraft flight manual supplement Regulatory and Guidance Library revolutions per minute Radio Technical Commission Of Aeronautics runway visual range Society of Automotive Engineers search and rescue stability augmentation system stability and control augmentation systems shaft horsepower stress vs. number of cycles structural repair manual system safety assessment supplemental type certificate short takeoff and landing time between overhaul type certificate type certificate data sheet takeoff decision point type inspection authorization type inspection report takeoff power turbine outlet temperature technical standard order true vapor pressure visual blade inspection method visual flight rules visual meteorological conditions ### Page 38 2/4/2016 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 HD HP VD VH VMO VNE VTOSS VX VY MMO NF NG NP NR CD CL CP CT AC 27-1B, Chg 7 Altitudes density altitude pressure altitude V speeds diving speed speed in level flight with maximum continuous power maximum operating limit speed never-exceed speed takeoff safety speed for Category A rotorcraft speed for best angle of climb speed for best rate of climb maximum operating mach number N speeds free turbine speed gas generator speed power turbine speed rotor speed Coefficients coefficient of drag coefficient of lift coefficient of power coefficient of thrust Page xxiii ### Page 39 **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 40 AC 21.31. § 21.31 TYPE DESIGN. The regulatory basis for requiring data to define 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 conforms 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 should 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 (Amendment 21-59) FLIGHT TESTS. a. Explanation. (1) This section outlines the requirements 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 41 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 tests 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 42 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 complexity 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 FAA/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 must assure that the FAA/AUTHORITY manufacturing inspector and certification engineer are aware of all configuration Page FAR 21 - 4 ### Page 43 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 responsibilities, 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 accomplished 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 rotorcraft 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 defects 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 single 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 44 AC 21.39. § 21.39 (Amendment 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 base 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 varies 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 45 **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 46 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 restraint 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 certificated 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 certified 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. (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, Page A - 2 ### Page 47 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, concerns 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 for 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 48 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 49 **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, that 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 minimum 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 50 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 helicopters, 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 conducted 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 certification 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 51 AC 27.25. § 27.25 (Amendment 27-14) WEIGHT LIMITS. a. Explanation. (1) This section is definitive and specifies 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 rotor 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 52 AC 27-1B, Chg 4 9/17/2009 appr oval. 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 demonstrate 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 increase 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, Certification 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 27.25A. § 27.25 (Amendment 27- 44) WEIGHT LIMITS. a. Explanation. Amendment 27-44 adds a requirement to create weight, altitude, and temperature limitations for those rotorcraft that are not able to meet the basic requirements of §§ 27.79 or 27.143(c). The § 27.79 height-velocity diagram must be demonstrated at maximum weight or at the highest weight allowing out-of-ground-effect (OGE) hover at a density altitude of 7,000 feet. The § 27.143(c) controllability must be demonstrated at critical weight at 7,000 feet density altitude. If either or both of these Page B – 4 ### Page 53 requirements cannot be met, the applicant must have in place the appropriate limitations to assure that take-offs and landings are limited to those weights, altitudes, and temperatures, plus any associated wind constraints, for which satisfactory demonstration has been made. In no case should those limits be established at an altitude that is not operationally suitable. In the past, the minimum operationally suitable altitude for takeoff and landing has been established as 3,000 feet density altitude. b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect. Page B – 4.1 ### Page 54 AC 27-1B, Chg 4 3/31/2014 AC 27.27.§ 27.27 (Amendment 27-2) CENTER OF GRAVITY LIMITS. a. Explanation. (1) This regulation is definitive and requires 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. (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 limits 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 the 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 Page B – 4.2 ### Page 55 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. (4) In summary, it is of prime importance 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. Page B – 5 ### Page 56 AC 27-1B, Chg 4 3/31/2014 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 (including both required and optional equipment) in the rotorcraft. It includes fixed ballast, unusable fuel, other unusable fluids, and full operating fluids. “Full operating fluids” such as oils used in an engine, auxiliary power unit, main and auxiliary gearboxes, and hydraulic systems are considered “closed fluid systems” typically filled to a “full mark” indicator level. Fluids necessary for the operation of nonpermanently installed equipment (i.e., carry-on equipment) are not considered part of the empty weight. (1) A ballast is fixed when made a permanent part of the rotorcraft as a means of controlling the empty weight center of gravity (CG). (2) Installed equipment is any FAA-approved equipment attached to the rotorcraft with hardware and, as a result, becomes an integral part of the rotorcraft. The installation or removal of such equipment must be recorded in the aircraft equipment list. Compliance with paragraph (b) of § 27.29 is accomplished by the use of an equipment list specifying 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 manufacturer and is normally made on a production rotorcraft rather than a prototype. If the manufacturer has been issued a production certificate and wishes to avoid weighing each production rotorcraft, the manufacturer may make a detailed proposal defining the procedures it would use to establish an empty weight and CG. When the proposal is approved, the manufacturer 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 rotorcraft, 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 FAA-H-8083-1 (Pilots Weight and Balance Handbook) for a sample weight and balance procedure. (3) The weight and balance should be recalculated 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). Page B – 6 ### Page 57 c. Ballast Loading and Type. (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) The two types of ballasts that may be used in loading are solid 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. not result. Page B – 6.1 ### Page 58 AC 27-1B, Chg 4 3/31/2014 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 Page B – 6.2 ### Page 59 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. (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 Page B – 7 ### Page 60 AC 27-1B, Chg 4 3/31/2014 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 61 **SUBPART B - FLIGHT** PERFORMANCE AC 27.45. § 27.45 (Amendment 27-14) PERFORMANCE - GENERAL. a. Explanation. (1) Section 27.45 of 14 CFR regulations lists the rules and standards under which the performance requirements are to be met. This guidance provides general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables that 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 an approved methodology. The guidance within this section is considered an approved methodology. (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 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. (4) The 30-second and 2-minute OEI power ratings, defined in 14 CFR 1.1, are based on up to three periods of use during a single flight. The purpose of the three periods is (i) to initially recover from an engine failure, (ii) to conduct OEI missed approach, and (iii) to conduct the final OEI landing. Rotorcraft performance based on the use of these time-limited power ratings is only permitted once in each of the above three periods (i.e., 30-second power must not be used more than once during the initial recovery from an engine failure). (5) All engines operating (AEO) performance must be based upon approved AEO power ratings. OEI power ratings cannot be applied to an AEO condition. b. Procedures. (1) Winds for Testing. Page B - 9 ### Page 62 AC 27-1B, Chg 4 3/31/2014 (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 accounted for, hover performance testing should be conducted in winds of 3 knots or less. (ii) IGE controllability 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 foregoing, 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) HV - 0 to 5 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: Extrapolation and Interpolation. (i) Using FAA//AUTHORITY approved methodology: (A) Hover performance may be extrapolated from test data up to a maximum of ±4,000 feet density altitude from the test altitude. (B) IGE handling qualities, height-velocity, and engine operating characteristics may be extrapolated from test data up to a maximum of ±2,000 feet density altitude from the test altitude. (C) Cruise stability and controllability tests should be evaluated at a minimum 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 density altitude. Page B - 10 ### Page 63 (ii) As in all testing, extrapolation 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. Explanation. 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. (A) 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 enroute altitude. (B) Maximum takeoff and landing altitude may be limited by the IGE hover ceiling for a rotorcraft as described in § 27.73 (through Amendment 27-14) or by low speed controllability. The hover ceiling and any information pertinent to takeoff and landing are presented in the performance information section of the RFM. For rotorcraft certificated to CAR 6, Amendment 6-7 or any amendment of part 27, a hover ceiling may not be presented above the altitude at which takeoff, HV, and IGE controllability tests were conducted plus allowable extrapolation, unless that extrapolated altitude is to at least 7,000 feet. If the applicant elects to demonstrate these tests to an altitude below 7,000 feet, including allowable extrapolation, 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) HV envelope (§ 27.79). (6) IGE controllability (§ 27.143(c)). (7) Cooling (§§ 27.1041, 27.1043, and 27.1045). Page B - 11 ### Page 64 AC 27-1B, Chg 4 3/31/2014 (8) Engine operating characteristics (§ 27.939). Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this guidance section. (B) As detailed in paragraph b.(2) above, the maximum allowable extrapolation of HV, 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. (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 landing 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 paragraph b.(3)(ii)(A) above. (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 materials. Therefore, to satisfy § 27.1309(a), the Page B - 12 ### Page 65 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 or interpolation of test data should only be allowed if the applicant’s predicted or calculated data is verified by actual test. Extrapolations or interpolations should not exceed 10°C below or 20°C above those values tested. (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 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, the engine manufacturer is responsible to ascertain that each engine delivered will produce, as a minimum, the certified power Page B - 13 ### Page 66 AC 27-1B, Chg 4 3/31/2014 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 installation losses (i.e., power decrements) 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 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 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 Page B - 14 ### Page 67 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 increase, engine temperature or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM 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 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 exceed a limit. These readings should be obtained approximately 15 seconds after power is initially applied. (2) To determine the 30-minute and 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. OEI 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. Page B - 15 ### Page 68 AC 27-1B, Chg 4 3/31/2014 (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 a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the RFM 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 may 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 Page B - 16 ### Page 69 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 or 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 engine bleed air pressure for their function. (3) The maximum attainable gas producer speed, and thus power available under certain ambients, may be affected if engine bleed air pressure is an input to the fuel scheduling mechanism. (4) Systems for surge protection which schedule on engine bleed air pressure 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 engine bleed air pressure 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 b.(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 b.(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. Page B - 17 ### Page 70 AC 27-1B, Chg 4 3/31/2014 (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 b.(5) above. 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. 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 RFM. b. Procedures. All of the guidance material pertaining to AC section 27.45 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 27-21) GENERAL. a. Explanation. Although § 27.45 was not changed by Amendment 27-29, that amendment added requirements for certification of 30-second/2-minute 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. MG 9 of this AC includes guidance material on power assurance procedures to assure that the OEI power level can be achieved. All of the guidance material pertaining to sections 27.45 and 27.45A of this AC remain in effect. AC 27.45C. § 27.45 (Amendment 27-21) PERFORMANCE - GENERAL. a. Explanation. Although § 27.45 was not changed by Amendment 27-44, that amendment added new performance and handling qualities requirements for normal category rotorcraft. Included within these regulatory changes is OGE handling qualities. Additionally, hover performance requirements were re-identified from § 27.73 to § 27.49; HV envelope requirements were re-identified from § 27.79 to § 27.87. Page B - 18 ### Page 71 b. Procedures. All of the guidance material pertaining to AC sections 27.45, 27.45A, and 27.45B remain in effect. In addition, the following apply: (1) OGE handling qualities may be extrapolated from test data up to a maximum of ±2,000 feet density altitude from the test altitude. (2) Hover performance guidance that applied to § 27.73 is applicable to § 27.49. (3) HV envelope guidance that applied to § 27.79 is applicable to § 27.87. Page B - 19 ### Page 72 AC 27-1B, Chg 4 3/31/2014 CERTIFICATION BASIS 14 CFR part 27 CAR 6 CAR 06 6-Amdt. Rqmts 27-Amdt. 21 27-Amdt. 2 Original 6-Amdt. 7 4 Original Original 1. MGW 1. MGW Sea 1. MGW 1. MGW 1. MGW 1. MGW 1. No specific Sea Level Level Sea Level Sea Level Sea Sea wt. and alt. H-V Ref. 2. Max. OGE 2. Wt. selected by applicant at 2. Wt. selected 2. Wt. selected Level 2. Wt. Level 2. Wt. requirements. 27.25 wt. Lesser lessor of: by by selected selected 27.79 27.1519 27.1587 Test Conditions of: a. Max. alt. cap. a. Max. alt. cap. b. 7000' Hd applicant at lessor of: applicant at lessor of: by applicant max. t.o. by applicant at 6.116 6.741 b. 7000' Hd a. Max. alt. cap. a. Max. alt. cap. and ldg. alt. max. t.o. and 6.743 b. 7000' b. 7000' ldg. alt. Hd Hd 3. H-V is 3. H-V is perf. 3. H-V is 3. H-V is 3. H-V is 3. H-V is H-V is perf. Info. 4. Max. alt. info. 4. If H-V wt is perf. info. perf. info. perf. info. operating limitation. operating limitation. for which less than IGE RFM H-V is valid. wt., wt. becomes limitation for t.o./ldg. AC 27-1B 5. If H-V is Section less than 27.45 & OGE wt., H- 27.51 V wt. becomes limit. 6. Applicant is encouraged to demo H-V Remarks 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 HV Requirements Page B - 20 ### Page 73 Rqmts 3/31/2014 AC 27-1B, Chg 4 CERT **IFICATION BASIS IGE** CONTROL Ref. 27.25 27.143 27.1587 AC 27-1B, Sections 27.45 & 27.143 Test Conditions RFM Remarks 27-Amdt. 21 1. MGW Sea Level 2. Max. IGE wt. Lesser o f: a. Max. alt. cap. b. 7000' Hd 3 . Critical CG 4. Critical Rotor RPM 5. Wi nd of not less t han 17 kts. 6. Max. safe wind is perf. inf o. 7. If 17 kts. wind demo'd to alt. less than 7000', a corresponding WAT limit must be established. 14 CFR part 27 Original 1. MGW Sea Level 2. Wt. selected by applicant to max. t.o. and ldg. alt. 3. Critical CG 4. Critical Rotor RPM 5. Cr itical wt. 6. Wind of not less t han 20 mph 7. Max. safe wind is perf. inf o. CAR 06 Original Original 1. MGW Sea 1. No specific Level requirement. 2. Max. approved wt. for t.o./ldg at alt above sea level. 3. Critical CG 4. Critical Rotor RPM 5. Wi nd not less than 20 mph. 6. Max. safe wind is perf. inf o. FIGURE AC 27.45-2 IGE CONTROLLABILITY REQUIREMENTS Page B - 21 ### Page 74 600 500 • 400 300 • 200 , ENG SIN 12345.-- ,'\\ ~ . f,' MIN CERTIFIED // __ SPECIFICATION E'Ni3IN~ 100 0 ·1000 1100 1200 1300 1400 ·1500 CORRECTED TOT°F FIGURE AC:27.45.-3 SHAF'T HOR~EPOWER VS"TURBlNE OUTLET TEMPERATURE --SEA LEVEL STANDARD DAV AC 27-1B, Chg 4 3/31/2014 Page B - 22 ### Page 75 0 iD 0 ~ ~ I- LLI w LL ' w 0 :::) I- I- _J ~ LLI 0:: :::) (fj fB O::'. (L t6, 14 IZ 111 6 SHAFf H0RSEP0YVER AVAILABLE FIGURE AC:2745-4 !JNINSTALLEDT AKEOFF POWER AVAILABLE Page B - 23 ### Page 76 AC 27-1B, Chg 4 3/31/2014 20 .I .I I I tjl I .I I I ..-1---+---1- •- -r-·- ·-r g>, I '.I I I i • I .I I I ii;l I I I I I• I ·1 I I ~-+--r--i----i- 1 ' I I ZI I I I I QI j I I ({)I I I I tl)lf- I I __ Tl __ Tl -~L .2:_1_ • :I _j I I I Z I I I I I I I I ! -+- +-- ---i- 1 I I I I I I I j I I I I . I -1-- -1---t- ! I I I I I I I j I I I -~--1---t I I I I I I I I I I I I '--L ---1. I ( I I I I I I I I -1--, I I I I I I I --, - I SHAFT HORSEPOWER .A.VAILABLE FIGURE AC27.45-5 INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM Page B - 24 ### Page 77 ~ C ::0 rTI :P Ci r,:; " f --0 0 ~ rn ::0 }> 01 (/) C s; z (") m 0 :r rn 0 ;,;; () ::t. }> ~ eoo 78il 760 14o 11o· !O"0 68!l 66!l ·c "C "" ·t c "C 'C 'C 3/31/2014 AC 27-1B, C hg 4 Pag e B – 24.1 ### Page 78 AC 27-1B, Chg 4 9/17/2009 AC 27.49. § 27.49 (Former § 27.73) (Amendment 27- 44) Performance at Minimum Operating Speed. (For § 27.73 prior to Amendment 27-44, see AC 27.73) a. Explanation. Amendment 27-44 adds a requirement to determine out-of-groundeffect (OGE) hover performance. Once reserved for special missions, OGE operations are now a common practice. (1) The word "hover" applies to a helicopter 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, § 27.49, refers to hover in–ground-effect (IGE) and OGE. Hover OGE is the absence of measurable ground effect. Hover OGE is established when the power required to hover is the same at different heights above the ground. (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 non-dimensional power coefficients (Cp) and thrust coefficients (Ct) for normalizing and presenting test results minimizes the amount of data required to cover the helicopter'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 IGE 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 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 - 24 2 ### Page 79 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. IGE hover performance may be affected by the composition and slope of the surface beneath the rotorcraft. Therefore, the tests should be conducted over a smooth, level, hard surface. (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 wind velocity should be less than 3 knots as there are no accurate methods of correcting hover data for wind effects. To minimize inaccuracies due to hysteresis, collective movement should be made in only one direction. Rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are: (i) Engine and transmission torque. (ii) Rotor speed. (iii) Ambient and engine temperatures, such as Measured Gas Temperature (MGT). (iv) Pressure altitude. (v) Fuel used (or remaining). (vi) Load cell reading. (vii) Generator(s) load. (viii) Wind speed and direction. (ix) Hover height. Page B – 24.3 ### Page 80 AC 27-1B, Chg 4 9/17/2009 As a technique, it is recommended the rotorcraft be loaded to a center of gravity (CG) near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spot-check using the free flight technique. The free flight technique in AC 27.49.b.(4) will determine if any problems, 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 equipment 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 re-ballasted 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. Applicants must account for any rotorcraft stability augmentation system that may influence hover performance. (5) Comprehensive hover performance tests are typically conducted at low, intermediate (~7000 feet Hd), and high altitude test sites, with prepared landing surfaces, in conjunction with takeoff, landing, controllability, and maneuverability testing. Alternatively, a predicted hover performance model developed for high altitude may be used if verified by limited flight-testing. The extrapolation guidelines in AC 27.45.b.(2) are still applicable. These higher altitude hover tests could typically be conducted in conjunction with the limited controllability tests. If the applicant is able to demonstrate to the FAA/AUTHORITY a method to provide a reliable hover reference, it is acceptable to conduct OGE tests without ground reference. Hover performance can usually be extrapolated up to a maximum of 4,000 feet above the highest test site altitude. Page B - 24 4 ### Page 81 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 should 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 wheels, 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 Page B -25 ### Page 82 AC 27-1B, Chg 4 9/17/2009 horizontal acceleration close to the ground has pilot scan outside the cockpit and 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 established 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 27.51A. § 27.51 (Amendment 27-44) Takeoff. a. Explanation. Amendment 27-44 revised the requirement to perform the test at the most critical center-of-gravity (CG) location as opposed to the most forward CG. While it is not always the case, the forward CG is particularly critical for most rotorcraft. Additionally, the change clarifies that the test must be performed at the maximum weight requested for takeoff for altitudes above sea level. The previous requirement stated that the test be performed at a weight selected by the applicant for altitudes above sea level. That weight was traditionally interpreted to mean the maximum requested takeoff weight for that altitude and it is now stated that way in the rule. b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect. Page B - 26 ### Page 83 [Page intentionally left blank.] Page B – 26.1 ### Page 84 AC 27-1B, Chg 4 3/31/2014 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 helicopters, 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 determined, the minimum climb gradient must be 1:6 for standard sea level conditions. (2) For helicopters, VY 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). [Section AC 27.65 continued on next page.] Page B – 26.2 ### Page 85 (3) For helicopters, if VNE 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 VY. (1) Sawtooth climbs may be used to determine VY. 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 power) 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 rotorcraft 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-Engine-Operating Climb Performance. Page B - 27 ### Page 86 (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. Page B - 28 ### Page 87 b. Procedures. (1) The procedure discussed in paragraph 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 accomplished 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 27-44) AUTOROTATION PERFORMANCE. All of the policy material pertaining to this section remains in effect. a. Explanation. (1) Performance capabilities during stabilized 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 importance, 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 autorotation 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 usually critical; however, center of gravity effects should be spot-checked to confirm this for a given design. Page B - 29 ### Page 88 b. Procedures. (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 89 Al RBF!EEIJ Fl 1.sURE AC 27 7'1- I AUr0RtJTATI 0NAL C:HAP·A~TERI-STleG -TYPICAL Pag e B - 31 ### Page 90 AC 27-1B, Chg 4 9/17/2009 AC 27.73. § 27.73 PERFORMANCE AT MINIMUM OPERATING SPEED (thru amendment 27-43; see § 27..49 effective with amendment 27-44). All of the policy material pertaining to AC 27.73 remains in effect through Amendment 27-43. See § 27.49 and AC 27.49 effective with Amendment 27-44. a. Explanation. (1) The word “hover” applies to a rotorcraft 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 Page B - 32 ### Page 91 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 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 wind 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. Page B - 33 ### Page 92 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, 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 equipment 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 difficult 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. Page B - 34 ### Page 93 AC 27.75. § 27.75 (Amendment 27-14) LANDING. a. Explanation. (1) This rule incorporates all of the landing requirements for Part 27 rotorcraft. (2) As with other flight maneuvers, 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 path 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 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. Aircraft 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 similar 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. Page B -34.1 ### Page 94 AC 27-1B, Chg 4 3/31/2014 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 gravity 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 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/σ (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 this 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 Page 34.2 ### Page 95 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, the 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 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 27.75A. § 27.75 (Amendment 27- 44) Landing. a. Explanation. Amendment 27-44 rewords § 27.75(a) by replacing the word “glide” with “autorotation” and further clarifies that the one-engine-inoperative (OEI) approach is to be performed from an established OEI approach. The OEI approach should be made utilizing a normal helicopter approach angle of approximately 6° or some other angle determined to be acceptable to the FAA/Authority and properly referenced in the Rotorcraft Flight Manual (RFM). b. Procedures. The policy material pertaining to this section remains in effect with the following changes and additions: Page B -35 ### Page 96 AC 27-1B, Chg 4 9/17/2009 (1) Instrumentation/Equipment. Aircraft 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 OEI landing is similar in many respects to the Height-Velocity (HV) tests described in AC 27.79 of this advisory circular. 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 OEI. The speed is reduced and the rotorcraft is flared to a conventional OEI landing. To show compliance, full OEI landing should be demonstrated from sea level to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less, at the maximum landing weight, without damage. For altitudes above 7,000 feet to the maximum takeoff and landing altitude, compliance with this requirement is shown by demonstrating that the OEI descent rate and forward speed can be reasonably controlled. (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 the 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 CG is usually the 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 of § 27.75(a)(1) and § 27.75(b). These paragraphs contain two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. The second 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. Typically a full autorotation landing to touchdown is demonstrated at sea level standard day for the maximum landing gross weight for that altitude. For altitudes above standard sea level, Page 36 ### Page 97 a demonstration of flare effectiveness with power recovery at the maximum landing gross weight corresponding to the altitude satisfies the requirement for touchdown. The flare must reduce the autorotative descent rate and forward speed to a reasonable value (i.e., autorotative descent rate and forward speed consistent with that demonstrated at sea level). If a complete company test program has documented the all-engine-out landing capability, verification tests may be initiated at those limiting weight conditions. If not, buildup testing should be initiated at light weight. (ii) Demonstrated compliance with this requirement is intended to show that the autorotative descent rate and forward speed at touchdown (less than 40 knots true airspeed (KTAS) is recommended) can be controlled to a reasonable value 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 application of collective pitch 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 application of collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude. Some designs incorporate features that 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 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, showing that the rate of descent and forward speed can be controlled to allow for safe landing. Other methods may be considered if they lead to reasonable assurance that rate of descent and forward speed can be 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 OEI 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, the 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 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 Page B - 36.1 ### Page 98 AC 27-1B, Chg 4 9/17/2009 to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as it is reasonable to dissipate speed and build RPM. Rotor RPM must stay within allowable limits. Aft CG is ordinarily critical due to visibility and pitch attitude effectiveness in a flare. 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. Page - 36.2 ### Page 99 AC 27.79. § 27.79 LIMITING HEIGHT-SPEED ENVELOPE. All of the policy material pertaining to this AC section remains in effect through Amendment 27-43. See § 27.87 and this AC’s section 27.87 effective with Amendment 27-44. 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 one engine inoperative (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 failure, 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 or 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 the 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, use takeoff power (or a lower power selected by the applicant as an operating procedure) and normal pilot reaction time for corrective control actuation. (5) Since the HV diagram may represent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer’s 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 Page B - 37 ### Page 100 AC 27-1B, Chg 4 3/31/2014 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 ground 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 lesser 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 (OGE). 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) The FAA accepts, as a method of extrapolation, a weight penalty of 3% for each 1000 feet above the permitted 2000 feet extrapolation. This weight penalty has been applied to extrapolations along a constant gross weight divided by density ratio, GW/σ. (10) Vertical takeoff and landing (VTOL) 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 101 (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 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 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 “flyin” 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 that comply with the category A engine isolation design requirements, the HV envelope may be determined with OEI 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 Page B - 39 ### Page 102 AC 27-1B, Chg 4 3/31/2014 failure tests. Care must be taken to ensure 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 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 103 (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 and 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 defined, 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 gear Page B - 41 ### Page 104 AC 27-1B, Chg 4 3/31/2014 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 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 105 FIGURE AC 27.79-1 HEIGHT-VELOCITY (HV) DIAGRAM Page B - 43 ### Page 106 AC 27-1B, Chg 4 3/31/2014 \ ~ 0 1 Appro:.:. ·~ lSOOlilt fs\(l[/;r Vi?lco::lly LA _J WeigJ1t CONSTANTH\/ DIAGRAM. VARIABLEWEIGHT Height ln~reasing DeflSitY Albtu rJ& Velot ity '-s__J - CONSTANT WEIGHT I ncrieasin.g GrossW~1ghf Vel:on1ty I_ C _J CONSTANT PEl~JSITY ALTITUDE FI.GIJRE AC,2J.7_Q.:2 ALT!lUD.ENILEJ~HT AC.CQ1Jt\l1ASIUTY Page B - 44 ### Page 107 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. AC 27.87. § 27.87 (Former § 27.79) (Amendment 27-44) Height-Speed Envelope. (For § 27.79 prior to Amendment 27-44, see AC 27.79) AC 27.87. § 27.87 (Amendment 27- 44) HEIGHT-SPEED ENVELOPE. a. Explanation. Amendment 27-44, in addition to some minor text changes, clarifies the one-engine-inoperative (OEI) engine power to be used (for multi-engine rotorcraft) when demonstrating the requirement. The engine power of the remaining engine is the minimum uninstalled specification power after it is corrected for installation losses. The methods of determining this power are established in the general performance paragraph of this AC. b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect. Page B - 45 ### Page 108 AC 27.87A. § 27.87 (AMENDMENT 27-51) HEIGHT-VELOCITY ENVELOPE. a. Explanation. Amendment 27-51 changed the title of the section from “Heightspeed envelope” to “Height-velocity envelope” to be consistent with the title nomenclature of § 29.87, which is the equivalent requirement for transport category rotorcraft. The title of § 29.87 was changed from “Limiting height-speed envelope” to “Height-velocity envelope” at amendment 29-39. b. Procedures. The policy pertaining to the procedures in this section remains unchanged. For purposes of this AC, the terms “height-velocity” and “height-speed” are equivalent. Page B - 45.1 ### Page 109 **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 must 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 Appendix B of this AC. (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 (RFM) 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 110 (4) During the FAA/AUTHORITY flight 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 assessment and to carefully document the results of these evaluations. Section 27.695 requires an alternate system allowing continued safe flight and landing following any single failure of a control system hydraulic boost system. This assessment of ‘safe’ should take into account not only residual post-failure control loads but also workload and pilot fatigue considerations. The following is suggested as an appropriate test sequence, conducted by a range of pilots, for a VFR approval: (i) Simulated hydraulic failure at critical flight conditions and Max GW. (ii) Establish level flight at a cruise speed >VY. (iii) Fly for approximately 30 minutes (to assess workload and account for pilot strength variations), demonstrating ability to climb or descend and small bank angle turns. This also allows for the possibility that the hydraulic failure occurs over water or other undesirable landing area. (iv) Land at a suitable area using a recommended landing technique, appropriate to the emergency. (5) Because control loads typically increase at higher altitudes, failure of the control hydraulic boost system should also be investigated during high altitude testing, including landing at 7000’ HD, at Max GW/σ. Where approval for any other type of operation is requested (e.g., IFR, category A), an appropriate test sequence must be proposed to the FAA/Authority. Section 27.141(b) provides the regulatory requirements for these strength and skill requirements, as well as a smooth transition capability between appropriate flight conditions. These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft. Flight characteristics and pilot workload should be evaluated in all expected flight conditions, including actual turbulence. (6) For night or IFR approval, § 27.141(c) requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this guidance. b. Procedures. None. Page B - 47 ### Page 111 AC 27-1B, Chg 4 3/31/2014 AC 27.143. § 27.143 (Amendment 27-21) CONTROLLABILITY AND MANEUVERABILITY. a. Explanation. (1) This regulation contains the basic 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. During the assessment carried out under § 27.143(a)(2)(v) for rotorcraft in glide (i.e., autorotation), in addition to controllability and maneuverability, it should be shown by flight testing that the directional stability characteristics are sufficient to allow the pilot to control the rotorcraft without undue attention to heading at the speeds for minimum rate of descent and best angle of glide. This should be evaluated at normal trim conditions and in turns up to 30° angle of bank. The ability to generate a sideslip must be evaluated throughout the autorotation speed envelope. 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.” The rotorcraft must still meet the stability requirements of Subpart B and, if applicable, Appendix B. (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 VNE 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. Nose-down pitching capability is needed for control of gust response and to allow necessary flight path changes in a nose-down 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 whether the remaining longitudinal control travel at VNE generate a clearly positive nose-down 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 (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 Page B – 48 ### Page 112 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) requires 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 that 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 (MCP) (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 the 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 MCP, 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. MCP 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’s feet being away from the pedals is 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 MCP, it does not limit engine failure testing to MCP. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions. Following power failure, Page B - 49 ### Page 113 AC 27-1B, Chg 4 3/31/2014 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 VNE (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 controllability 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 MCP 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 height above the terrain must be specified in the limitation section of the rotorcraft flight manual (RFM). (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), Page B – 50 ### Page 114 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 or softover malfunction should be included in the Rotorcraft Flight Manual. b. Procedures. (1) Flight test instrumentation should 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 and CG variation as fuel is burned. (2) The critical condition for VNE controllability testing is ordinarily aft CG, MCP, 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 MCP 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 V NE controllability test is normally 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 or temperature envelope should be analyzed and investigated by flight test. (3) The critical condition for controllability testing in a hover is ordinarily forward CG at maximum weight with minimum power-on rotor RPM. For rearward flight testing of Page B - 51 ### Page 115 AC 27-1B, Chg 4 3/31/2014 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. Testing will be carried out at the power required to achieve stabilized flight conditions. With rotorcraft that are operating in conditions such that the gross weight is limited by the power available, there should always be adequate tail rotor pedal authority to maintain yaw control when using the maximum approved all engines operating (AEO) power, which is take-off power for most designs. (i) Where the rotorcraft is capable of operating at maximum gross weight with less than maximum approved power, it is appropriate to examine the rotorcraft characteristics with small amounts of additional power applied above the trim power required to allow for typical power variations experienced during normal use of the rotorcraft. For example, maneuvering, turbulence, or rotor governing characteristics may cause the pilot to use power in excess of power required for trim. The maximum power excursions should not exceed maximum approved AEO power, excluding any transient range. (ii) The rotorcraft should be flown both in ground effect (IGE) and out of ground effect (OGE), with the most adverse wind speed and direction for directional control within the flight envelope proposed, using power variations above trim that might be expected during normal use of the rotorcraft. Consideration should be given to the amount of excess power available, the ease with which power can be controlled via the collective, the effect of tail rotor control inputs on power required, and the characteristics of the rotorcraft if the limits of directional control are approached. There should be no tendency to deviate rapidly or suddenly in yaw. This assessment is normally conducted in conjunction with the critical azimuth testing. (4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of the 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) Page B – 52 ### Page 116 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 VNE with any power setting must permit suitable nose-down 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 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 (see section 27.75 of this AC). Page B - 52.1 ### Page 117 AC 27-1B, Chg 4 9/17/2009 AC 27.143A. § 27.143 (Amendment 27- 44) Controllability and Maneuverability. a. Explanation. Amendment 27-44 made a minor clarification to assure that inground-effect (IGE) controllability is demonstrated at all speeds up to 17 knots. In many rotorcraft, the entry into the regime of translational lift requires the most power, thus potentially causing control difficulties, and frequently occurs at speeds less than 17 knots. The amendment also requires that, above 7,000 feet density altitude in which takeoff and landing performance is scheduled, the controllability of the rotorcraft be determined. The amendment also requires that out-of-ground-effect (OGE) controllability be determined up to a speed of at least 17 knots at a weight selected by the applicant up to the maximum takeoff and landing altitude of the rotorcraft. All the policy material pertaining to this section remains in effect with the following changes: (1) This regulation contains the basic controllability requirements for normal category rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for controllability and for maneuverability are summarized in § 27.143(a), which is 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) in § 27.143 include more specific flight conditions and highlight the typical areas of concern during a flight test program. (i) § 27.143(b) specifies flight at VNE with critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at VNE for nose down pitching of the rotorcraft and for adequate roll control. Nose down pitching capability is needed for control of gust response and to allow necessary flight path changes in a nose down 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 nose down 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 (directionally) should be investigated. This "control remaining" philosophy must also be applied for other flight conditions specified in this section. (ii) § 27.143(c) requires a minimum control capability for hover and takeoff in winds from zero to at least 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 Page 52.2 ### Page 118 near the ground such as rolling takeoffs for wheeled rotorcraft. On helicopters 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) § 27.143(e) requires 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 that 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 (MCP) (or takeoff power where appropriate) including reasonable engine torque splits. For multiengine Category A installations with three or more engines, the 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 MCP 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. MCP 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 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 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 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 MCP, it does not limit engine failure testing to MCP. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions. Following power failure, the rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits. Page B – 52.3 ### Page 119 AC 27-1B, Chg 4 9/17/2009 (iv) § 27.143(f) addresses the special case in which a VNE (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 controllability 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) and § 27.175(c), and to 1.1 VNE (power-off) in autorotation per § 27.143(f) 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 height above the terrain must be specified in the limitation section of the Rotorcraft Flight Manual (RFM). (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. Page 52.4 ### Page 120 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 or softover malfunction should be included in the RFM. b. Procedures. The policy material pertaining to this section remains in effect with the following changes and additions: (1) Flight test instrumentation should 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 VNE 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 MCP 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 V NE controllability test is normally 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) Controllability. (i) The critical condition for controllability testing in a hover is ordinarily forward CG at maximum weight with minimum power-on rotor RPM. For rearward flight Page B – 52.5 ### Page 121 AC 27-1B, Chg 4 9/17/2009 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. IGE hover controllability testing should be accomplished with the lowest portion of the helicopter at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. OGE testing should be done with the rotor at a predetermined height above the ground at which it has been determined that there is no ground effect. 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 helicopter is held in the most critical position relative to wind. (A) Testing will normally be carried out at the power required to achieve stabilized flight conditions. However, it is also important to show that yaw control remains adequate to allow normal power changes that might be required in normal operational maneuvers typical for the type and use of the rotorcraft. With rotorcraft that are operating in conditions in which the gross weight is limited by the power available, there should always be adequate tail rotor pedal control available to maintain yaw control when using up to take-off power. However, this will not be the case if the rotorcraft weight in the low speed flight envelope is limited by yaw control system capability. (B) To cover the case where excess power is available, it is appropriate to examine the rotorcraft characteristics with some small amounts of additional power applied above the trim power required to hover to allow for typical power variations that will be experienced during normal use of the rotorcraft. For example, maneuvering or turbulence will cause the pilot to use some of the excess power available. The rotorcraft should be flown, both IGE and OGE, with the most adverse wind speed and direction for directional control within the flight envelope proposed. Use power variations above trim that might be expected during normal use of the rotorcraft giving consideration to the amount of excess power available, the ease with which power can be controlled by collective, and the characteristics of the rotorcraft if the limits of directional control are approached. There should be no tendency to deviate rapidly or suddenly in yaw. This assessment is normally conducted in conjunction with the critical azimuth testing. (C) It may be appropriate to provide flight manual information on the directional control characteristics, including any relevant maximum power above which it could be expected that directional control might not be maintained. Page 52.6 ### Page 122 (ii) Comprehensive controllability tests are typically conducted at low, intermediate (~7000 feet Hd) and high tests sites, with prepared landing surfaces, in conjunction with takeoff, landing, and performance testing. (iii) Alternatively, a predicted controllability model developed for high altitude may be used if verified by limited flight-testing with steady ambient winds. The extrapolation guidelines in AC 27.45 b (2) are still applicable. These high altitude controllability tests could typically be conducted in conjunction with takeoff, landing, and performance tests. (iv) Controllability can usually be extrapolated up to a maximum of 2,000 feet above the highest test site altitude. Note: Engine operating characteristics must be considered during the limited high altitude tests. (4) Prior to engine failure testing, the pilot should 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 type inspection authorization (TIA). Particular attention should be given to minimum stabilized and minimum transient rotor RPM limits. These values should 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 should 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 exceeding a temperature limitation. This is needed for compliance with § 29.903(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 VNE with any power setting must permit suitable nose down 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 that requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft 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 pitch-up, is satisfactory). Obviously, during the Page B – 52.7 ### Page 123 AC 27-1B, Chg 4 3/31/2014 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 pilotinduced 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 rotorcraft. 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. Care must 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 rotorcraft, 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 rotorcraft. Characteristics of the flight control system should be qualitatively considered during other flight tests. These characteristics include forces, control harmony, nonlinearities, discontinuities, proper directional senses, breakout, friction, hysteresis, etc. Page B – 52.8 ### Page 124 AC 27.161. § 27.161 (Amendment 27-21) TRIM CONTROL. a. Explanation. (1) The pilot has many tasks to perform 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 require 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 test 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 friction 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 125 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 which 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 careful 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. [See new AC 27.173A (dated 9/17/2009) posted as separate document in RGL with this Master AC.] a. Explanation. (1) This rule contains control requirements 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, through 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 126 b. Procedures. (1) The control requirement of this section 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 27.173A. § 27.173 (Amendment 27- 44) STATIC LONGITUDINAL STABILITY. a. Explanation. (1) Amendment 27-44 makes a major change to the requirement by allowing for neutral or negative static longitudinal stability in limited flight domains. Additionally the requirement for the hover demonstration found in § 27.173(c) has been deleted as this requirement is adequately covered by the controllability requirements. The basic tenants of the rule are unchanged in that the rule contains control system design requirements 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. This requirement in no way assures aircraft stability. It is simply a control requirement that speaks to direction of control motion. Rotorcraft with either highly stable or highly unstable static longitudinal stability characteristics can typically comply with the basic requirement for control sense of motion. All the policy material pertaining to this section remains in effect with the following changes and additions: (2) §§ 27.173 through 27.175 contain the basic control position requirements necessary to establish a minimum level of static longitudinal stability. Positive stability is found for conditions of climb, cruise, VNE, and autorotation in § 27.175 by demonstrating a stable stick position gradient through a specified speed range. This is the primary method of demonstrating compliance with the longitudinal static stability requirements. (3) For aircraft that do not possess positive control position stability for some limited flight conditions or modes of operation, an equivalent level of safety was previously provided which requires a qualitative evaluation of the pilot’s ability to maintain a given airspeed, within 5 knots of the desired speed, without exceptional piloting skill or alertness. These flight conditions and modes of operation could include various combinations of gross weight, CG, flight regime (climb, cruise, descent), ambient conditions (altitude/temperature) as well as possible variations in the stability augmentation configuration. In the past, the FAA/AUTHORITIES have certified numerous rotorcraft, under equivalent level of safety findings, which have neutral or negative static longitudinal stick position stability in some flight domains. This amendment to § 27.173 is intended to allow for this case without having to resort to an Page B – 55 ### Page 127 AC 27.1B, Chg4 9/17/2009 equivalent safety finding. For these previous equivalent safety findings, acceptable qualitative flight characteristics were found on aircraft, which possessed negative longitudinal stick position gradients of up to 2-3% of total control travel in certain flight regimes; however, this value is not intended to be a limit. When this means of compliance is elected by the applicant, in addition to the qualitative pilot evaluation, it is still necessary to collect the data associated with the classical static longitudinal stability testing as defined in § 27.175. b. Procedures. All the policy material pertaining to this section remains in effect with the following changes and additions: (1) The control requirement of paragraph (a) of this section 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 assure compliance with the primary stability requirements of this section are contained under § 27.175, Demonstration of Static Longitudinal Stability. Refer to AC 27.175 of this advisory circular for an explanation of detailed flight test procedures. (3) The procedures necessary to assure compliance with the alternative (i.e., pilot evaluation) method of compliance are provided below. (i) For those limited conditions where compliance with the basic control position requirements cannot be shown, the evaluation must focus on the ability of the pilot to maintain airspeed in the flight regime without exceptional piloting skill or alertness under typical flight conditions. “Limited flight conditions” infers that the aircraft should be in reasonable compliance with the stick position stability requirements of § 27.173(b) for most of the flight conditions and configurations tested. Extraordinary means of complying with § 27.173(b) should not be forced on the aircraft design if the airspeed retention task meets the pilot skill and alertness guidelines. The demonstration flight regimes are defined in § 27.175(a) through (d). For those flight regimes, conditions and configurations where compliance with stick position requirements of § 27.173(b) cannot be shown, the evaluation pilot should assess the ease of maintaining airspeed within the specified +/- 5 knots. (ii) When assessing the ease of maintaining airspeed the total workload must be considered. Secondary tasks pertinent to the minimum flight crew in each flight regime should be conducted. This may include visual navigation and communication in cruise, traffic avoidance in climb, and landing site selection in autorotation. (iii) The cues that the aircraft provides are an important contributor to the evaluation, and the nature of these cues should be noted in the compliance report where this alternate qualitative evaluation determines that the aircraft has satisfactory Page B – 56 ### Page 128 airspeed stability characteristics. The cues that supplant the control position cues may be found to be sufficient if these cues are natural to the speed maintenance task, and provide adequate guidance to the pilot during the task. One important cue might be the pitch attitude gradient with speed, where a perceptible change in trimmed pitch attitude is required for a perceptible airspeed change. Where pitch attitude is the predominant cue the relationship should be positive (nose down with airspeed increase) and perceptible without exceptional alertness. With this relationship, the evaluation pilot may find that the natural pitch control tasks associated with attitude control result in adequate airspeed retention, and the aircraft would be found to be in compliance. It may be that the power/airspeed relationship of the aircraft can create adequate cues, where a significant rate of descent is created by a nose down pitch attitude change and a subsequent airspeed increase. In this case, the normal cues associated with altitude retention during fixed power cruise flight may prove to be acceptable for airspeed retention if the evaluation pilot finds that, within the context of the overall flight task, airspeed retention is sufficiently accurate. These altitude change cues may not be usable in autorotation or climb, but may be sufficient in cruise or VNE tasks. (iv) Other cues may be found for a specific aircraft, such as small but perceptible changes in noise or vibration. It is not intended that the evaluation pilot search for these cues in order to learn how to maintain airspeed in the aircraft under evaluation. These cues should be perceptible to the typical pilot and sufficient to reinforce the airspeed maintenance task. Page B – 57 ### Page 129 AC 27.1B, Chg4 3/31/2014 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 instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when 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 parameters include pressure altitude, ambient temperature, and indicated airspeed (pace vehicle or theodolite speed for hover tests). Page B – 58 ### Page 130 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 VH (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 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 Page B – 59 ### Page 131 AC 27.1B, Chg4 3/31/2014 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 autorotation 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 verified 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 – 60 ### Page 132 90 ~ < ::i ,80 2 ~ ~ 10 LL 10 g O 1:RUISE D QIMB ~◊ P.UTORQT ATlOl'\1 ' hOVER 3HADEO POINTS REPRESEI\IT TRIM 0< e e ◄ ,'IRla,t power~of,f rriax.1111un1 des.i:gn IJmit rotor sp~-:e:d. 5 Forward f119ht 1. 11' VNE at 11ilininilltn' PQWE;lraQn rotor sps.etl. 6 Forward fli9ht 1.11' VNe at maximum p¢wer-on rotorsi;1eet1. 7. Rtgllt and left turn at VNE at rnatimum power-on rotor speed with 30' bank angle