# AC 29-2C CHG 9: Certification of Transport Category Rotorcraft

- Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it)
- Status: Current
- Approved: June 23, 2023
- Family: AC 29-2
- Pages: 1453
- FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/DRSDOCID174453317620230413220544.0001
- Page: https://openairdocs.org/ac/29-2

## Versions

- AC 29-2C CHG 9 (Current), approved Jun 23, 2023
- AC 29-2C CHG 4-8 (Historical), approved Jul 2, 2018
- AC 29-2C CHG 1-3 (Historical), approved Sep 30, 2008
- AC 29-2C CHG 2 (Historical), approved Apr 25, 2006
- AC 29-2C (Historical), approved Jun 24, 2005
- AC 29-2C (Historical), approved Apr 25, 2003
- AC 29-2C (Historical), approved Apr 25, 2003
- AC 29-2C (Historical), approved Apr 25, 2003
- AC 29-2C CHG 1 (Historical), approved Feb 12, 2003
- AC 29-2C (Historical), approved Feb 12, 2003
- AC 29-2B (Historical), approved Jul 30, 1997
- AC 29-2 CHG 1 (Historical), approved Nov 19, 1984
- AC 29-2C (Historical), approved 

## Text

### 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

0 AdvisoryU.S. Department of Transportation Federal Aviation Circular Administration Subject: Certification of Transport Date: 6/23/23 AC No: 29-2C Category 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: 29.955,

29.997, 29.1019, 29.1305, 29.1309, 29.1329, 29.1333, 29.1335, 29.1353, 29.1517, 29.1543,

29.1545, 29.1549, 29.1555, 29.1587, and AC 29 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 Digitally signed by

VICTOR W WICKLUND Date: 2023.06.23 16:30:39 -04'00' 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 Transport Category Date: 07/02/2018 AC No: 29-2C 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

29.773.

PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated D-68 9/30/1999 D-68.1, D-68.2 07/02/18

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

C Advisory US Department orTransportation CircularFederal Aviation Administration Subject: CERTIFICATION OF TRANSPORT Date: FEB 4 2016 AC No: 29-2C CATEGORY ROTORCRAFT Initiated by: ASW-110 Change: 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: 29.865B,

29.1303, 29.1321, 29 MG 1, 29 MG 16, and 29 Appendix B.

b. This change adds to this AC the following MG sections: AC 29 MG 19 (Guidance on Electronic Display Systems (EDS) for Rotorcraft Installations), AC 29 MG 20 (Human Factors (HF)), and AC 29 MG 21 (Guidance on Creating a System Level Functional Hazard Assessment (FHA)).

3. WEBSITE AVAILABILITY. To access this AC electronically, go to the Advisory Circulars library at http://www.faa.gov/regulations policies/advisory circulars/.

anager, Rotorcraft Directorate, Aircraft Certification Service

### Page 5

U.S. Deportment of Transportation Federal Aviation Administration Advisory Circular

Subject: CERTIFICATION OF TRANSPORT Date: 7/25/2014 AC No: 29-2C 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 29 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 29 Rotorcraft., 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/ Lance T. Gant, for Kim Smith Manager, Rotorcraft Directorate, Aircraft Certification Service

### Page 6

U.S. Department of Transportation Federal Aviation Administration Advisory Circular

Subject: CERTIFICATION OF TRANSPORT Date: 6/13/2014 AC No: 29-2C 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 transport 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

U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF TRANSPORT Date: 5/1/2014 AC No: 29-2C 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: 29.25A, 29.49, 29.143A, 29.173A, 29.175A, 29.177A, 29.1587B, and 29 Appendix B.

b. The changed AC 29 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 29.573 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

### Page 8

AC 29-2C, Chg 4 5/1/2014 Fatigue Evaluation of Composite Rotorcraft Structures rulemaking, is incorporated in this change 4.

d. The AC 29.571B section previous approved on December 2, 2011 and posted separately to the Regulatory and Guidance Library (RGL), which is related to the Fatigue tolerance Evaluation of Metallic Structure rulemaking, is incorporated in this change 4.

e. The changed AC 29.972 and 29.927A sections (Additional Tests for Rotor Drive Systems), previously approved on July 6, 2012 and posted separately to the Regulatory and Guidance Library (RGL), are incorporated in this change 4.

f. The changed AC 29 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.

g. The changed AC 29 MG 18 section on (Helicopter Terrain Awareness and Warning System (HTAWS)), previously approved on February 27, 2014, is incorporated in this change 4.

h. This change 4 also includes changes, edits, and additions for these AC sections: 29.29,

29.45, 29.49, 29.79, 29.141, 29.143, 29.151, 29.251, 29.337, 29.561, 29.562, 29.563A, 29.571A,

29.603, 29.610, 29.625, 29.801, 29.855A, 29.863, 29.865, 29.865B, 29.871, 29.903B, 29.923B,

29.939, 29.1011, 29.1093, 29.1093A, 29.1193, 29.1309, 29.1316, 29.1317, 29.1329, 29.1333,

29.1337, 29.1351, 29.1357, 29.1401, 29.1411, 29.1435. 29.1501, 29.1527, 29.1541, 29.1543,

29.1549, 29.1583, 29 MG 1, 29 MG 2, 29 MG 4, 29 MG 8, 29 MG 13, 29 MG 18, 29 MG 22, and 29 Appendix A.

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 2

### Page 9

US Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF TRANSPORT Date: 9/30/2008 AC No:29-2C 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 9 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 29.571, 29.679, 29.695, 29.783, 29.901A, 29.917A,

29.1307, 29.1351, and 29.1431 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 29-2C, Chg 3 9/30/2008 AC 29-2C, Chg 3

PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated vii thru x...........................2/12/03 & 4/25/06 xiii & xiv.......................... 2/12/03 & 4/25/06 C-93 thru C-96 ................. 9/30/99 & 2/12/03 D-41 thru D-44................. 9/30/99 & 4/25/06 D-47 & D-48 ..................................... 9/30/99 D-79 thru D-82.................................. 9/30/99 E-7 & E-8.......................................... 9/30/99 E-35 thru 38 ..................... 9/30/99 & 4/25/06 F-11 & F-12 ..................... 9/30/99 & 2/12/03 F-65 & F-66 ...................................... 9/30/99 F-113 thru F-116 ............................... 9/30/99 vii thru x .............2/12/03, 4/25/06, & 9/30/08 xiii thru xiv ......................................... 9/30/08 C-93 thru C-96 ................... 2/12/03 & 9/30/08 D-41 thru D-44 ................................... 9/30/08 D-47 thru D-48 .................. 9/30/99 & 9/30/08 D-79 thru D-82 ................................... 9/30/08 E-7 & E-8 .......................... 9/30/99 & 9/30/08 E-35 thru E-38 ................... 9/30/99 & 9/30/08 F-11 & F-12........................................ 9/30/08 F-65 thru F-66.2 ................................. 9/30/08 F-113 thru F-116 ............... 9/30/99 & 9/30/08 Signed by Scott A. Horn for Mark R. Schilling Acting Manager, Rotorcraft Directorate Aircraft Certification Service Page 

### Page 11

U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF TRANSPORT

CATEGORY ROTORCRAFT Date:

Initiated 4/25/06 ASW-110 AC No: 29-2C Change: Chg 2 ______________________________________________

**4. 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 25 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 that contains changed text. 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.

**5. PRINCIPAL CHANGES.**

a. Paragraphs 29.45, 29.59A, 29.75A, 29.337, 29.351B, 29.547A, 29.602, 29.631, 29.672, 29.683,

29.777, 29.917A, 29.1321, 29.1333, 29.1351, 29.1585, MG 1, MG 4, MG 8, MG16, Appendix A, and Appendix B are revised.

b. The AC material in paragraphs 29.801 and 29.1411 has been revised.

c. The AC material in MG 12 has been revised and is now contained in AC 29.865B, Subpart D.

d. New paragraph MG 18, Helicopter Terrain Awareness Warning System (HTAWS), is added to Chapter 3.

e. New figures AC 29.351A-1, 29.351B-1, 29.351B-2, and 29.865B-1 are added in Chapter 2.

f. New figures AC 29 MG 8-1, 8-2, 8-3, 8-4, 18-1, 18-2, and 18-2.1 are added in Chapter 3.

6. WEBSITE AVAILABILITY. To access this AC electronically, log on to http://www.airweb.faa.gov/rgl and then click on AC’s.

### Page 12

AC 29-2C, Chg 2 4/25/06 AC 29-2C, Chg 2

PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated iii thru x…………………….9/30/99 & 2/12/03 xiii thru xxiv……………….9/30/99 & 2/12/03 B-11 thru B-20 ………………………..9/30/99 B-49 thru B-125………………………..9/30/99 C-5 thru C-6 ………………………..9/30/99 C-13 thru C-96 …………….9/30/99 & 2/12/03 D-1 thru D-159 …………….9/30/99 & 2/12/03 E-37 thru E-48 ………………………..9/30/99 F-43 thru F-44 .. ………………………..9/30/99 F-61 thru F-120 …………….9/30/99 & 2/12/03 G-45 thru G-52 ………………………..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 12-44…...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 & 2/12/03 iii thru x...............................................4/25/06 xiii thru xxv......................................... 4/25/06 B-11 thru B-20 .................................... 4/25/06 B-49 thru B-127 .................................. 4/25/06 C-5 thru C-6 ........................................ 4/25/06 C-13 thru C-98 .................................... 4/25/06 D-1 thru D-185 .................................... 4/25/06 E-37 thru E-48 .................................... 4/25/06 F-43 thru F-44 ..................................... 4/25/06 F-61 thru F-123 ................................... 4/25/06 G-45 thru G-52 .................................... 4/25/06 MG 1-15 thru MG 1-16 ...................... 4/25/06 MG 4-9 thru MG 4-13 ........................ 4/25/06 MG 8-1 thru MG 8-31 ........................ 4/25/06 MG 12-1.............................................. 4/25/06 MG 16-1 thru MG 16-18 .................... 4/25/06 MG 17-1.............................................. 4/25/06 MG 18-1 thru MG 18-27 .................... 4/25/06 Apdx A-47 thru Apdx A-71................ 4/25/06 Apdx B-11 thru Apdx B-21 ................ 4/25/06 David A. Downey Manager, Rotorcraft Directorate Aircraft Certification Service Page 

### Page 13

U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: CERTIFICATION OF TRANSPORT

CATEGORY ROTORCRAFT Date:

Initiated 2/12/03 ASW-110 AC No: 29-2C Change: Chg 1 ______________________________________________

**7. PURPOSE.**

a. This change revises existing material in 21 paragraphs and adds new material for seven paragraphs.

b. The change number and the date of the changed material are shown at the top of each page that contains changed text. 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.

**8. CANCELLATION.**

a. AC 20-95, Fatigue Evaluation of Rotorcraft Structure, May 18, 1976, is canceled in its entirety.

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 29.562.

c. AC test containing references to AC 20-95 and AC 20-137 was either changed or deleted in paragraphs AC 29.2, 29.562, 29.613, 29.785A, 29.865A, 29.907, 29.952, MG 8, and MG 12.

**9. PRINCIPAL CHANGES.**

a. Paragraphs 29.562, 29.571, 29.805, 29.1303, 29.1309, 29.1353, 29.1419, 29.1505,

29.1529, MG 5, MG 6, MG 11, MG 12, and Appendix B are revised.

b. New paragraphs 29.625A and 29.785B are added to Chapter 2.

c. New paragraphs MG 13, MG 14, MG 15, and MG 16 are added to Chapter 3.

d. Appendix A (previously reserved) is added.

e. New figures AC 29.562-1, 29.562-2, 29.562-3, 29.562-4, 29.562-5, 29.562-6, 29.562-7,

29.562-8, 29.1309-1, 29.1309-2, 29.1309-3, 29.1309-4, and 29.1309-5 are added in Chapter 2.

f. New figures AC 29 MG 11-2, 11-4, 11-5, 14-1, 15-1, and 15-2 are added in Chapter 3.

10. WEBSITE AVAILABILITY. To access this AC electronically, log on to http://www.airweb.faa.gov/rgl and then click on AC’s.

### Page 14

AC 29-2C, Chg 1

PAGE CONTROL CHART Remove Pages Dated Insert Pages Dated v thru x .............................................. 9/30/99 xiii thru xxiii....................................... 9/30/99 A-3 thru A-4 ................................. 9/30/99 C-51 thru C-58 ................................. 9/30/99 D-13 thru D-14 ................................. 9/30/99 D-27 thru D-30 ................................. 9/30/99 D-85 thru D-100 ................................ 9/30/99 D-133 thru D-134 .............................. 9/30/99 E-29 thru E-30 . ................................. 9/30/99 E-57 thru E-58 ................................. 9/30/99 E-69 thru E-70 ................................. 9/30/99 E-75 thru E-80 ................................. 9/30/99 F-1 thru F-105 ................................. 9/30/99 G-3 thru G-50 ................................. 9/30/99 MG 5-1 thru MG 5-6 ......................... 9/30/99 MG 6-5 thru MG 6-18 ....................... 9/30/99 MG 8-5 thru MG 8-6 ......................... 9/30/99 MG 11-1 thru MG 11-25 ................... 9/30/99 MG 12-23 thru MG 12-24 ................. 9/30/99 MG 12-37 thru MG 12-38 ................. 9/30/99 MG 12-43 thru MG 12-44 ................. 9/30/99 none none none none none Apdx B-13 thru Apdx B-14 ............... 9/30/99 v thru x ............................................... 2/12/03 xiii thru xxiv....................................... 2/12/03 A-3 thru A-4....................................... 2/12/03 C-51 thru C-96 ................................... 2/12/03 D-13 thru D-14 ................................... 2/12/03 D-27 thru D-30 ................................... 2/12/03 D-85 thru D-100 ................................. 2/12/03 D-133 thru D-134 ............................... 2/12/03 E-29 thru E-30 .................................... 2/12/03 E-57 thru E-58 .................................... 2/12/03 E-69 thru E-70 .................................... 2/12/03 E-75 thru E-80 .................................... 2/12/03 F-1 thru F-120 .................................... 2/12/03 G-3 thru G-52 ..................................... 2/12/03 MG 5-1 thru MG 5-6 .......................... 2/12/03 MG 6-5 thru MG 6-20 ........................ 2/12/03 MG 8-5 thru MG 8-6 .......................... 2/12/03 MG 11-1 thru MG 11-22.................... 2/12/03 MG 12-23 thru MG 12-24.................. 2/12/03 MG 12-37 thru MG 12-38.................. 2/12/03 MG 12-43 thru MG 12-44.................. 2/12/03 MG 13-1 thru 13-6 ............................. 2/12/03 MG 14-1 thru 14-8 ............................. 2/12/03 MG 15-1 thru 15-18 ........................... 2/12/03 MG 16-1 thru 16-16 ........................... 2/12/03 Apdx A-1 thru Apdx A-70 ................. 2/12/03 Apdx B-13 thru Apdx B-14 ............... 2/12/03 David A. Downey Manager, Rotorcraft Directorate Aircraft Certification Service 2

### Page 15

Advisory Circular _____________________________________________ U.S. Department of Transportation Federal Aviation Administration Subject: CERTIFICATION OF TRANSPORT Date: 9/30/99 AC No: 29-2C CATEGORY ROTORCRAFT Initiated by: ASW-110 Change:

1. PURPOSE:

a. This is a total revision of AC 29-2B 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 29-2C and replaces AC 29-2B in its entirety. This revises existing material in 12 paragraphs, adds new material for two 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 29, which contains the Airworthiness Standards for Transport Category Rotorcraft. It includes methods of compliance in the areas of basic design, ground tests, and flight tests.

2. CANCELLATION. AC 29 -2B, Certification of Transport 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), Transport Category Rotorcraft," with the following changes:

• Paragraphs that correspond to a FAR number are merged into existing AC text in Chapter 2.

• Paragraphs that do not correspond with a FAR number either remain in Chapter 3 and are renumbered as MG paragraphs, or are now an appendix.

• In order to stay aligned with FAR numbering, Appendices A, C, and D are reserved for future AC material.

### Page 16

b. Paragraphs revised to incorporate technical guidance are AC 21.35, 29.561, 29.610, 29.863,

29.952, 29.1093, 29.1305, 29.1309, 29.1351, 29.1353, and MG 4 (FADEC). Paragraph AC

29.661, Rotorblade Clearance, contains new material added as a result of National Transportation Safety Board (NTSB) recommendations.

c. New paragraphs added are 29.602, Critical Parts, and MG 12, External Loads. These paragraphs correspond with recent harmonized regulatory changes. Also, figure 29.863- 1 is new.

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 rotorc raft 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 29. 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 29. 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 29 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 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 29. Each paragraph has the applicable amendment to Part 29 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, which follow the original numbered paragraphs. These subsequent paragraphs provide updated guidance information or changes to policy that parallel a specific rule change.

The guidance material in the original paragraph (for earlier amendments) still applies and is modified as explained in each of the later paragraphs for later amendments. The applicable amendment number will only appear in the title line for the “A,” “B,” etc., paragraphs. The guidance material in the initial paragraph is intended to apply to all amendments except as modified by the later paragraphs. Each ensuing “A,” “B,” etc., paragraph will be identified with an amendment level to indicate the rule change that precipitated the policy change.

8. RELATED PUBLICATIONS. FAA Certification personnel and designees should be familiar with Order 8110.4, Type Certification, and Order 8100.5, Aircraft Certification Directorate Procedures.

Eric Bries Acting Manager, Rotorcraft Directorate 2

### Page 17

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 29 AIRWORTHINESS STANDARDS:

**TRANSPORT CATEGORY ROTORCRAFT**

**SUBPART A - GENERAL**

**GENERAL**

AC 29.1 Applicability … A-1

AC 29.1A Applicability … A-2

AC 29.2 Special Retroactive Requirements … A-3

**SUBPART B - FLIGHT**

**GENERAL**

AC 29.21 Proof of Compliance … B-1

AC 29.21A Proof of Compliance … B-3

AC 29.25 Weight Limits … B-3

AC 29.25A Weight Limits … B-4

AC 29.27 Center of Gravity Limits … B-5

AC 29.29 Empty Weight and Corresponding Center of Gravity … B-6

AC 29.31 Removable Ballast … B-8

AC 29.33 Main Rotor Speed and Pitch Limits … B-8

**PERFORMANCE**

AC 29.45 Performance - General … B-10

AC 29.45A Performance - General … B-20

AC 29.45B Performance - General … B-20

AC 29.49 Performance at Minimum Operating Speed...................................... B-26.2 

### Page 18

AC 29-2C, Chg 5 6/13/2014 Page No.

SUBPART B - FLIGHT (continued) PERFORMANCE (continued)

AC 29.51 Takeoff Data - General … B-29

AC 29.51A Takeoff Data - General … B-30

AC 29.53 Takeoff: Category A … B-31

AC 29.53A Takeoff: Category A … B-34

AC 29.55 Takeoff Decision Point: Category A … B-36

AC 29.59 Takeoff Path: Category A … B-36

AC 29.59A Takeoff Path, Distance and Rejected Takeoff, Ground Level and ... B-46 Elevated Heliport: Category A

AC 29.63 Takeoff: Category B … B-57

AC 29.65 Climb: (All Engines Operating) … B-60

AC 29.65A Climb: (General and All Engines Operating) … B-61

AC 29.67 Climb: One Engine Inoperative … B-63

AC 29.67A Climb: One Engine Inoperative … B-65

AC 29.71 Rotorcraft Angle of Glide: Category B … B-66

AC 29.73 Performance at Minimum Operating Speed … B-69

AC 29.75 Landing … B-71

AC 29.75A Landing … B-80

AC 29.77 Balked Landing: Category A … B-91

AC 29.79 Limiting Height - Speed Envelope … B-92

AC 29.85 Balked Landing: Category A … B-100

AC 29.87 Limiting Height - Speed Envelope … B-101

**FLIGHT CHARACTERISTICS**

AC 29.141 Flight Characteristics – General … B-102

AC 29.143 Controllability and Maneuverability … B-104

AC 29.143A Controllability and Maneuverability … B-106

AC 29.151 Flight Controls … B-108

AC 29.161 Trim Control … B-109

AC 29.161A Trim Control … B-110

AC 29.171 Stability: General … B-110

AC 29.173 Static Longitudinal Stability … B-111

AC 29.173A Static Longitudinal Stability … B-111

AC 29.175 Demonstration of Static Longitudinal Stability................................... B-112.2

AC 29.175A Demonstration of Static Longitudinal Stability … B-116

AC 29.177 Static Directional Stability.................................................................. B- AC 29.177A Static Directional Stability.................................................................. B-116.5

AC 29.181 Dynamic Stability: Category A Rotorcraft … B-117

Page 

### Page 19

Page No.

SUBPART B - FLIGHT (continued)

**GROUND AND WATER HANDLING CHARACTERISTICS**

AC 29.231 General (Ground and Water Handling Characteristics) … B-118

AC 29.235 Taxiing Condition … B-118

AC 29.239 Spray Characteristics … B-119

AC 29.241 Ground Resonance … B-120

**MISCELLANEOUS FLIGHT REQUIREMENTS**

AC 29.251 Vibration … B-125

**SUBPART C - STRENGTH REQUIREMENTS**

**GENERAL**

AC 29.301 Loads … C-1

AC 29.303 Factor of Safety … C-1

AC 29.305 Strength and Deformation … C-2

AC 29.307 Proof of Structure … C-2

AC 29.307A Proof of Structure … C-3

AC 29.309 Design Limitations … C-4

**FLIGHT LOADS**

AC 29.321 General - Flight Loads … C-5

AC 29.337 Limit Maneuvering Load Factor … C-5

AC 29.339 Resultant Limit Maneuvering Loads … C-8

AC 29.341 Gust Loads … C-8

AC 29.351 Yawing Conditions … C-10

AC 29.351A Yawing Conditions … C-13

AC 29.351B Yawing Conditions … C-15

AC 29.361 Engine Torque … C-18

**CONTROL SURFACE AND SYSTEM LOADS**

AC 29.391 Control Surface and System Loads - General … C-20

AC 29.391A Control Surface and System Loads - General … C-20

AC 29.395 Control System … C-20

AC 29.395A Control System … C-22

AC 29.397 Limit Pilot Forces and Torques … C-22

AC 29.399 Dual Control System … C-23

AC 29.401 Auxiliary Rotor Assemblies … C-23

Page 

### Page 20

AC 29-2C, Chg 5 6/13/2014 Page No.

SUBPART C - STRENGTH REQUIREMENTS (continued) CONTROL SURFACE AND SYSTEM LOADS (continued)

AC 29.401A Auxiliary Rotor Assemblies … C-24

AC 29.403 Auxiliary Rotor Attachment Structure … C-25

AC 29.403A Auxiliary Rotor Attachment Structure … C-25

AC 29.411 Ground Clearance: Tail Rotor Guard … C-26

AC 29.413 Stabilizing and Control Surfaces … C-27

AC 29.413A Stabilizing and Control Surfaces … C-29

AC 29.427 Unsymmetrical Loads … C-29

**GROUND LOADS**

AC 29.471 Ground Loads - General … C-33

AC 29.473 Ground Loading Conditions and Assumptions … C-33

AC 29.475 Tires and Shock Absorbers … C-34

AC 29.477 Landing Gear Arrangement … C-34

AC 29.479 Level Landing Conditions … C-34

AC 29.481 Tail-Down Landing Conditions … C-36

AC 29.483 One-Wheel Landing Conditions … C-36

AC 29.485 Lateral Drift Landing Conditions … C-36

AC 29.493 Braked Roll Conditions … C-37

AC 29.497 Ground Loading Conditions: Landing Gear With Tail Wheels … C-37

AC 29.501 Ground Loading Conditions: Landing Gear With Skids … C-39

AC 29.501A Ground Loading Conditions: Landing Gear With Skids … C-40

AC 29.505 Ski Landing Conditions … C-41

AC 29.511 Ground Load: Unsymmetrical Loads on Multiple- Wheel Units … C-41

**WATER LOADS**

AC 29.519 Hull Type Rotorcraft: Water -Based, Amphibian … C-43

AC 29.521 Float Landing Conditions … C-45

**MAIN COMPONENT REQUIREMENTS**

AC 29.547 Main Rotor Structure … C-46

AC 29.547A Main Rotor and Tail Rotor Structure … C-46

AC 29.549 Fuselage and Rotor Pylon … C-50

AC 29.551 Auxiliary Lifting Surfaces … C-50

Page 

### Page 21

Page No.

SUBPART C - STRENGTH REQUIREMENTS (continued)

**EMERGENCY LANDING CONDITIONS**

AC 29.561 Emergency Landing Conditions - General … C-51

AC 29.561A Emergency Landing Conditions - General … C-53

AC 29.561B Emergency Landing Conditions - General … C-54

AC 29.562 Emergency Landing Dynamic Conditions … C-55

AC 29.563 Structural Ditching Provisions … C-92

AC 29.563A Structural Ditching Provisions … C-92

**FATIGUE EVALUATION**

AC 29.571 Fatigue Evaluation of Flight Structure … C-94

AC 29.571A Fatigue Evaluation of Flight Structure … C-97

AC 29.571B Fatigue Tolerance Evaluation of Metallic Structure … C-98

AC 29.573 Damage Tolerance and Fatigue Evaluation of Composite … C-130

Rotorcraft Structures

**SUBPART D - DESIGN AND CONSTRUCTION**

**GENERAL**

AC 29.601 Design … D-1

AC 29.602 Critical Parts … D-2

AC 29.603 Materials … D-4

AC 29.605 Fabrication Methods … D-5

AC 29.607 Fasteners … D-5

AC 29.609 Protection of Structure … D-6

AC 29.610 Lightning and Static Electricity Protection … D-8

AC 29.611 Inspection Provisions … D-12

AC 29.613 Material Strength Properties and Design Values … D-13

AC 29.613A Material Strength Properties and Design Values … D-14

AC 29.619 Special Factors … D-17

AC 29.621 Casting Factors … D-17

AC 29.623 Bearing Factors … D-27

AC 29.625 Fitting Factors ................................................................................... D-28.1 AC 29.625A Fitting Factors .................................................................................. D-28.1

AC 29.629 Flutter … D-29

AC 29.629A Flutter and Divergence … D-29

AC 29.631 Bird Strike … D-31

Page 

### Page 22

AC 29-2C, Chg 5 6/13/2014 Page No.

SUBPART D - DESIGN AND CONSTRUCTION (continued) ROTORS

AC 29.653 Pressure Venting and Drainage of Rotor Blades … D-32

AC 29.659 Mass Balance … D-32

AC 29.661 Rotor Blade Clearance … D-32

AC 29.663 Ground Resonance Prevention Means … D-34

AC 29.663A Ground Resonance Prevention Means … D-36

**CONTROL SYSTEMS**

AC 29.671 Control Systems - General … D-37

AC 29.672 Stability Augmentation, Automatic, and Power-Operated Systems... D-37

AC 29.673 Primary Flight Controls … D-38

AC 29.674 Interconnected Controls … D-39

AC 29.675 Stops … D-39

AC 29.679 Control System Locks … D-41

AC 29.681 Limit Load Static Tests … D-42

AC 29.683 Operation Tests … D-42

AC 29.685 Control System Details … D-43

AC 29.687 Spring Devices … D-45

AC 29.691 Autorotation Control Mechanism … D-45

AC 29.695 Power Boost and Power-Operated Control System … D-46

**LANDING GEAR**

AC 29.723 Shock Absorption Tests … D-49

AC 29.725 Limit Drop Test … D-52

AC 29.727 Reserve Energy Absorption Drop Test … D-52

AC 29.727A Reserve Energy Absorption Drop Test … D-53

AC 29.729 Retracting Mechanism, Landing Gear … D-53

AC 29.731 Wheels … D-55

AC 29.733 Tires … D-56

AC 29.735 Brakes … D-58

AC 29.737 Skis … D-59

**FLOATS AND HULLS**

AC 29.751 Main Float Buoyancy … D-61

AC 29.753 Main Float Design … D-61

AC 29.755 Hull Buoyancy … D-64

AC 29.757 Hull and Auxiliary Float Strength … D-65

Page 

### Page 23

Page No.

SUBPART D - DESIGN AND CONSTRUCTION (continued)

**PERSONNEL AND CARGO ACCOMMODATIONS**

AC 29.771 Pilot Compartment … D-67

AC 29.773 Pilot Compartment View … D-68

AC 29.775 Windshields and Windows … D-74

AC 29.775A Windshields and Windows … D-74

AC 29.775B Windshields and Windows … D-74

AC 29.777 Cockpit Controls … D-75

AC 29.779 Motion and Effect of Cockpit Controls … D-76

AC 29.783 Doors … D-78

AC 29.783A Doors … D-82

AC 29.785 Seats, Safety Belts, and Harnesses … D-83

AC 29.785A Seats, Berths, Belts, Safety Belts, and Harnesses … D-85

AC 29.785B Seats, Berths, Litters, Safety Belts, and Harnesses … D-87

AC 29.787 Cargo and Baggage Compartments … D-88

AC 29.787A Cargo and Baggage Compartments … D-88

AC 29.801 Ditching … D-89

AC 29.803 Emergency Evacuation … D-97

AC 29.803A Emergency Evacuation … D-97

AC 29.805 Flightcrew Emergency Exits … D-98

AC 29.805A Flightcrew Emergency Exits … D-98

AC 29.807 Passenger Emergency Exits … D-99

AC 29.807A Emergency Exits … D-100

AC 29.809 Emergency Exit Arrangement … D-101

AC 29.809A Emergency Exit Arrangement … D-102

AC 29.811 Emergency Exit Marking … D-103

AC 29.811A Emergency Exit Marking … D-105

AC 29.812 Emergency Lighting … D-106

AC 29.813 Emergency Exit Access … D-106

AC 29.815 Main Aisle Width … D-108

AC 29.831 Ventilation … D-109

AC 29.833 Heaters … D-110

**FIRE PROTECTION**

AC 29.851 Fire Extinguishers … D-112

AC 29.853 Compartment Interiors … D-113

AC 29.855 Cargo and Baggage Compartments … D-115

AC 29.855A Cargo and Baggage Compartments … D-117

AC 29.859 Combustion Heater Fire Protection … D-117

AC 29.861 Fire Protection of Structure, Controls, and Other Parts … D-121

AC 29.861A Fire Protection of Structure, Controls, and Other Parts … D-124

Page ix

### Page 24

Page No.

SUBPART D - DESIGN AND CONSTRUCTION (continued) FIRE PROTECTION (continued)

AC 29.863 Flammable Fluid Fire Protection … D-124

**EXTERNAL LOAD ATTACHING MEANS**

AC 29.865 External Load Attaching Means … D-129

AC 29.865A External Load Attaching Means … D-134

AC 29.865B External Load Attaching Means (SEE MG 12) … D-140

**MISCELLANEOUS (DESIGN AND CONSTRUCTION)**

AC 29.871 Leveling Marks … D-165

AC 29.873 Ballast Provisions … D-165

AC 29.877 Ice Protection … D-166

**SUBPART E - POWERPLANT**

**GENERAL**

AC 29.901 Installation … E-1

AC 29.901A Installation … E-8

GENERAL (continued)

AC 29.901B Installation … E-8

AC 29.903 Engines … E-9

AC 29.903A Engines … E-11

AC 29.903B Engines … E-12

AC 29.903C Turbine Engine Installation … E-13

AC 29.907 Engine Vibration … E-29

AC 29.908 Cooling Fans … E-29

AC 29.908A Cooling Fans … E-30

**ROTOR DRIVE SYSTEM**

AC 29.917 Design … E-33

AC 29.917A Design … E-34

AC 29.921 Rotor Brake....................................................................................... E-37.2

AC 29.923 Rotor Drive System and Control Mechanism Tes ts … E-39

AC 29.923A Rotor Drive System and Control Mechanism Tests … E-45

AC 29.923B Rotor Drive System and Control Mechanism Tests … E-52

Page x

### Page 25

6/23/23 AC 29.923C AC 29.927 AC 29.927A AC 29.931 AC 29.935 AC 29.939 AC 29.951 AC 29.952 AC 29.953 AC 29.954 AC 29.955 AC 29.955A AC 29.955B AC 29.957 AC 29.959 AC 29.961 AC 29.961A AC 29.963 AC 29.963A AC 29.963B AC 29.965 AC 29.967 AC 29.967A AC 29.969 AC 29.969A AC 29.971 AC 29.971A AC 29.973 AC 29.973A AC 29.975 AC 29.975A AC 29.975B AC 29.977 AC 29.979 AC 29-2C, Chg 9 Page No.

SUBPART E - POWERPLANT (continued) ROTOR DRIVE SYSTEM (continued)

Rotor Drive System and Control Mechanism Tests … E-53

Additional Tests ................................................................................ E-54.1

Additional Tests … E-56

Shafting Critical Speeds … E-58

Shafting Joints … E-63

Turbine Engine Operating Characteristics … E-63

**FUEL SYSTEM**

Fuel System - General … E-66

Fuel System Crash Resistance … E-68

Fuel System Independence … E-93

Fuel System Lightning Protection … E-94

Fuel Flow … E-95

Fuel Flow … E-99

Fuel Flow .......................................................................................... E-102.1 I

Flow Between Interconnected Tanks … E-103

Unusable Fuel Supply … E-103

Fuel System Hot Weather Operation … E-104

Fuel System Hot Weather Operation … E-106

Fuel Tanks: General … E-107

Fuel Tanks: General … E-107

Fuel Tanks: General … E-108

Fuel Tank Tests … E-109

Fuel Tank Installation … E-111

Fuel Tank Installation … E-113

Fuel Tank Expansion Space … E-113

Fuel Tank Expansion Space … E-114

Fuel Tank Sump … E-114

Fuel Tank Sump … E-116

Fuel Tank Filler Connection … E-116

Fuel Tank Filler Connection … E-117

Fuel Tank Vents and Carburetor Vapor Vents … E-118

Fuel Tank Vents and Carburetor Vapor Vents … E-119

Fuel Tank Vents and Carburetor Vapor Vents … E-119

Fuel Tank Outlet … E-120

Pressure Refueling and Fueling Provisions Below Fuel Level … E-121

Page 

### Page 26

AC 29-2C, Chg 9 6/23/23 Page No.

SUBPART E - POWERPLANT (continued)

**FUEL SYSTEM COMPONENTS**

AC 29.991 Fuel Pumps … E-123

AC 29.991A Fuel Pumps … E-124

AC 29.993 Fuel System Lines and Fittings … E-124

AC 29.995 Fuel Valves … E-125

AC 29.997 Fuel Strainer or Filter … E-125

AC 29.997A Fuel Strainer or Filter … E-126

AC 29.997B Fuel Strainer or Filter ........................................................................ E-127.1 AC 29.999 Fuel System Drains........................................................................... E-127.2 AC 29.999A Fuel System Drains........................................................................... E-127.2

AC 29.1001 Fuel Jettisoning … E-128

**OIL SYSTEM**

AC 29.1011 Oil System - General … E-130

AC 29.1019B Oil Strainer or Filter........................................................................... E-134.1 AC 29.1021 Oil System Drains ............................................................................. E-134.2 AC 29.1025 Oil Valves.......................................................................................... E-134.2

AC 29.1013 Oil Tanks … E-131

AC 29.1015 Oil Tank Tests … E-131

AC 29.1017 Oil Lines and Fittings … E-132

AC 29.1019 Oil Strainer or Filter … E-132

AC 29.1019A Oil Strainer or Filter … E-133

AC 29.1023 Oil Radiators … E-134

AC 29.1027 Transmission and Gearboxes: General … E-135

**COOLING**

AC 29.1041 Cooling - General … E-137

AC 29.1043 Cooling Tests … E-140

AC 29.1043A Cooling Tests … E-142

AC 29.1045 Cooling Test Procedures … E-143

AC 29.1047 Takeoff Cooling Test Procedures … E-147

AC 29.1049 Hovering Cooling Test Procedures … E-149

**INDUCTION SYSTEM**

AC 29.1091 Air Induction … E-151

AC 29.1093 Induction System Icing Protection … E-152

AC 29.1093A Induction System Icing Protection ..................................................... E-160.2

AC 29.1101 Carburetor Air Preheater Design … E-161

AC 29.1103 Induction System Ducts and Air Duct Systems … E-161

AC 29.1105 Induction System Screens … E-164

Page 

### Page 27

Page No.

SUBPART E - POWERPLANT (continued) INDUCTION SYSTEM (continued)

AC 29.1107 Intercoolers and After -Coolers … E-165

AC 29.1109 Carburetor Air Cooling … E-165

**EXHAUST SYSTEM**

AC 29.1121 General … E-166

AC 29.1123 Exhaust Piping … E-167

AC 29.1125 Exhaust Heat Exchangers … E-169

**POWERPLANT CONTROLS AND ACCESSORIES**

AC 29.1141 Powerplant Controls: General … E-173

AC 29.1142 Auxiliary Power Unit Controls … E-174

AC 29.1143 Engine Controls … E-175

AC 29.1143A Engine Controls … E-176

AC 29.1143B Engine Controls … E-176

AC 29.1145 Ignition Switches … E-176

AC 29.1147 Mixture Controls … E-177

AC 29.1151 Rotor Brake Controls … E-178

AC 29.1157 Carburetor Air Temperature Controls … E-179

AC 29.1159 Supercharger Controls … E-179

AC 29.1163 Powerplant Accessories … E-180

AC 29.1165 Engine Ignition Systems … E-181

**POWERPLANT FIRE PROTECTION**

AC 29.1181 Designated Fire Zones:  Regions Included … E-182

AC 29.1183 Lines, Fittings, and Components … E-183

AC 29.1185 Flammable Fluids … E-184

AC 29.1187 Drainage and Ventilation of Fire Zones … E-186

AC 29.1189 Shutoff Means … E-186

AC 29.1191 Firewalls … E-187

AC 29.1193 Cowling and Engine Compartment Covering … E-191

AC 29.1193A Cowling and Engine Compartment Covering … E-192

AC 29.1194 Other Surfaces … E-193

AC 29.1195 Fire Extinguishing Systems … E-194

AC 29.1197 Fire Extinguishing Agents … E-197

AC 29.1199 Extinguishing Agent Containers … E-200

AC 29.1201 Fire Extinguishing System Materials … E-201

AC 29.1203 Fire Detector Systems … E-201

Page xiii

### Page 28

6/23/23AC 29-2C, Chg 9 Page No.

**SUBPART F - EQUIPMENT**

**GENERAL**

AC 29.1301 Function and Installation … F-1

AC 29.1305D Powerplant Instruments .................................................................... F-12.1 AC 29.1307 Miscellaneous Equipment ................................................................. F-12.3

AC 29.1303 Flight and Navigation Instruments … F-1

AC 29.1305 Powerplant Instruments … F-7

AC 29.1305A Powerplant Instruments … F-8

AC 29.1305B Powerplant Instruments … F-9

AC 29.1305C Powerplant Instruments … F-12

AC 29.1309 Equipment, Systems, and Installations … F-13

AC 29.1309A Equipment, systems, and Installations … F-40

AC 29.1309B Equipment, systems, and Installations.............................................. F-40.1

AC 29.1316 Electrical and Electronic System Lightning Protection … F-41

AC 29.1317 High-Intensity Radiated Fields (HIRF) Protection … F-42

**INSTRUMENTS - INSTALLATION**

AC 29.1321 Arrangement and Visibility … F-43

AC 29.1322 Warning, Caution, and Advisory Lights ............................................. F-44.2 AC 29.1329A Automatic Pilot System. .................................................................... F-60.1 AC 29.1329B Automatic Pilot and Flight Guidance System. ................................... F-60.2a AC 29.1331 Instruments Using a Power Supply ................................................... F-60.3 AC 29.1333A Instrument Systems .......................................................................... F-61.1 AC 29.1335A Flight Director Systems ..................................................................... F-63.1 AC 29.1337 Powerplant Instruments - Fuel Gauge Calibration............................ F-63.2

AC 29.1323 Airspeed Indicating System … F-48

AC 29.1323A Airspeed Indicating System … F-49

AC 29.1325 Static Pressure Systems … F-50

AC 29.1327 Magnetic Direction Indicator … F-51

AC 29.1329 Automatic Pilot Systems … F-51

AC 29.1333 Instrument Systems … F-61

AC 29.1335 Flight Director Systems … F-62

AC 29.1337A Powerplant Instruments … F-64

**ELECTRICAL SYSTEMS AND EQUIPMENT**

AC 29.1351 Electrical Systems and Equipment - General … F-66

AC 29.1353A Energy Storage Systems................................................................... F-78.1 AC 29.1359 Electrical System Fire and Smoke Protection ................................... F-82.1

AC 29.1353 Electrical Equipment and Installations … F-74

AC 29.1355 Distribution System … F-79

AC 29.1355A Distribution System … F-79

AC 29.1357 Circuit Protective Devices … F-80

AC 29.1357A Circuit Protective Devices … F-82

AC 29.1363 Electrical Systems Tests … F-83

Page xiv

### Page 29

AC 29-2C, Chg 5 Page No.

SUBPART F - EQUIPMENT (continued) LIGHTS

AC 29.1381 Instrument Lights … F-84

AC 29.1383 Landing Lights … F-85

AC 29.1385 Position Light System Installation … F-85

AC 29.1387 Position Light System Dihedral Angles … F-85

AC 29.1389 Position Light Distribution and Intensities … F-85

AC 29.1391 Minimum Intensities in the Horizontal Plane of Forward and Rear

Position Light … F-85

AC 29.1393 Minimum Intensities in Any Vertical Plane of Forward and Rear

Position Lights … F-85

AC 29.1395 Maximum Intensities In Overlapping Beams of Forward and Rear

Position Lights … F-85

AC 29.1397 Color Specifications … F-85

AC 29.1399 Riding Light … F-86

AC 29.1401 Anticollision Light System … F-87

**SAFETY EQUIPMENT**

AC 29.1411 Safety Equipment - General … F-88

AC 29.1413 Safety Belts: Passenger Warning Device … F-90

AC 29.1415 Ditching Equipment … F-91

AC 29.1419 Ice Protection … F-93

**MISCELLANEOUS EQUIPMENT**

AC 29.1431 Electronic Equipment … F-114

AC 29.1433 Vacuum Systems … F-114

AC 29.1435 Hydraulic Systems … F-115

AC 29.1439 Protective Breathing Equipment … F-118

AC 29.1457 Cockpit Voice Recorder … F-119

AC 29.1459 Flight Recorders … F-120

AC 29.1461 Equipment Containing High Energy Rotors … F-123

**SUBPART G - OPERATING LIMITATIONS AND INFORMATION**

AC 29.1501 Operating Limitations - General … G-1

**OPERATING LIMITATIONS**

AC 29.1503 Airspeed Limitations: General … G-1

AC 29.1505 Never-Exceed Speed … G-3

Page 

### Page 30

AC 29-2C, Chg 9 6/23/23 Page No.

SUBPART G - OPERATING LIMITATIONS AND INFORMATION (continued) OPERATING LIMITATIONS (continued)

AC 29.1509 Rotor Speed … G-6

AC 29.1517A Limiting Height-Velocity Envelope..................................................... G-9.1 AC 29.1519 Weight and Center of Gravity............................................................ G-9.2 AC 29.1521 Powerplant Limitations...................................................................... G-9.2

AC 29.1517 Limiting Height-Speed Envelope … G-8

AC 29.1521A Powerplant Limitations … G-10

AC 29.1521B Powerplant Limitations … G-11

AC 29.1522 Auxiliary Power Unit Limitations … G-12

AC 29.1523 Minimum Flightcrew … G-12

AC 29.1525 Kinds of Operations … G-16

AC 29.1527 Maximum Operating Altitude … G-17

AC 29.1529 Instructions for Continued Airworthiness (Maintenance Manual) … G-17

AC 29.1529A Instructions for Continued Airworthiness (Maintenance Manual) … G-22

**MARKINGS AND PLACARDS**

AC 29.1541 Marking and Placards - General … G-23

AC 29.1545 Airspeed Indicator ............................................................................. G-27.6 AC 29.1545A Airspeed Indicator ............................................................................. G-27.7 AC 29.1547 Magnetic Direction Indicator.............................................................. G-27.9 AC 29.1549 Powerplant Instruments .................................................................... G-27.9 AC 29.1549A Powerplant Instruments .................................................................... G-27.9

AC 29.1543 Instrument Markings: General … G-23

AC 29.1543A Instrument Markings: General.......................................................... G-27.1 AC 29.1549B Powerplant Instruments .................................................................... G-27.10

AC 29.1551 Oil Quantity Indicators … G-28

AC 29.1553 Fuel Quantity Indicator … G-28

AC 29.1555 Control Markings … G-29

AC 29.1555A Control Markings............................................................................... G-29.2

AC 29.1557 Miscellaneous Markings and Placards … G-30

AC 29.1559 Limitations Placard … G-31

AC 29.1561 Safety Equipment … G-32

AC 29.1565 Tail Rotor … G-33

**ROTORCRAFT FLIGHT MANUAL**

AC 29.1581 Rotorcraft Flight Manual - General … G-34

AC 29.1583 Operating Limitations … G-42

AC 29.1583A Operating Limitations … G-45

AC 29.1585 Operating Procedures … G-45

AC 29.1587 Performance Information … G-47

AC 29.1587A Performance Information … G-51

AC 29.1587B Performance Information … G-51

AC 29.1589 Loading Information AC 29.1587C Performance Information.................................................................. G-54.1

. … G-51

Page 

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Page No.

**CHAPTER 3 - MISCELLANEOUS GUIDANCE (MG),**

TRANSPORT CATEGORY ROTORCRAFT AC 29 MG 1 Certification Procedure for Rotorcraft Avionics Equipment............. MG 1-1 System Tests Interior Arrangements, and Equipment Flotation System Rotorcraft Metallic Structure (Including Flaw Tolerance) Air Conditioning Systems Systems (HUMS) Systems (NVIS) Equipment System AC 29 MG 18 Helicopter Terrain Awareness and Warning System (HTAWS) ...... MG 18-1 Installations Assessment (FHA) Installation in Part 29 Rotorcraft AC 29 MG 2 Standardized Test Procedure for Rotorcraft DC Electrical ............. MG 2-1 AC 29 MG 3 Rotorcraft and Systems Certification for Category II Operations .... MG 3-1 AC 29 MG 4 Full Authority Digital Electronic Controls (FADEC) ......................... MG 4-1 AC 29 MG 5 Agricultural Dispensing Equipment Installation ............................... MG 5-1 AC 29 MG 6 Emergency Medical Service (EMS) Systems, Installations, .......... MG 6-1 AC 29 MG 7 Structural Condition Indicators ....................................................... MG 7-1 AC 29 MG 8 Substantiation of Composite Rotorcraft Structure .......................... MG 8-1 AC 29 MG 9 Rotorcraft One- Engine-Inoperative (OEI) Power Assurance .......... MG 9-1 AC 29 MG 10 Advisory Material for Substantiation of Emergency ........................ MG 10-1 AC 29 MG 11 Fatigue Tolerance Evaluation of Transport Category ..................... MG 11-1 AC 29 MG 12 External Loads ............................................................................... MG 12-1 AC 29 MG 13 Systems Certification Considerations ............................................ MG 13-1 AC 29 MG 14 Certification Procedure for Installation of Vapor Cycle .................. MG 14-1 AC 29 MG 15 Airworthiness Approval of Rotorcraft Health Usage Monitoring ..... MG 15-1 AC 29 MG 16 Certification Procedure for Rotorcraft Night Vision Imaging .......... MG 16-1 AC 29 MG 17 Guidance on Analyzing an Advanced Flight Controls (AdFC) ........ MG 17-1 AC 29 MG 19 Guidance on Electronic Display Systems (EDS) for Rotor craft ...... MG 19-1 AC 29 MG 20 Human Factors (HF) ....................................................................... MG 20-1 AC 29 MG 21 Guidance on Creating a System Level Functional Hazard ............. MG 21-1 AC 29 MG 22 Rotorcraft One Engine Inoperative (OEI) Training Mode................ MG 22-1 AC 29 MG 23 Automatic Flight Guidance and Control Systems (AFGCS) ............ MG 23-1 Page xvii

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AC 29-2C, Chg 9 6/23/23

APPENDICES AC 29 Appendix A – Instructions for Continued Airworthiness.................................. Apdx A-1 AC 29 Appendix B - Airworthiness Guidance for Rotorcraft Instrument Flight........... Apdx B-1 AC 29 Appendix C - (reserved) (Icing Certification).................................................. Apdx C-1 AC 29 Appendix D - (reserved) (Criteria for Demonstration of Emergency )............ Apdx D-1 Evacuation Procedures Under § 29.803 Page xviii

### Page 33

FIGURES The first digits in the figure number are the same as the related paragraph number. The digits after the hyphen are the number within the paragraph. Most figures are located at the end of the related paragraph.

Figure No. Title Page No.

AC 29.45-1 H-V Requirements. B-22 AC 29.45-2 IGE Controllability Requirements. B-23 AC 29.45-3 Shaft Horsepower vs. Turbine Outlet Temperature - B- 24 Sea Level Standard Day.

AC 29.45-4 Uninstalled Takeoff Power Available. B-25 AC 29.45-5 Installed Takeoff Power Available, Anti-Ice Off, 400 RPM. B-26 AC 29.45-6 Power Assurance Check Chart. B-26.1 AC 29.53-1 Takeoff Performance Category A. B-33 AC 29.53A-1 Takeoff Performance Category A. B-35 AC 29.59-1 Category A Vertical Takeoff Profile Ground Level Heliport. B-44 AC 29.59-2 Category A Vertical Takeoff Profile Pinnacle. B-45 AC 29.59A-1 Category A Vertical Takeoff Profile Ground Level Heliport. B-55 AC 29.59A-2 Category A Vertical Takeoff Profile Elevated Heliport. B-56 AC 29.63-1 Conventional Takeoff Profile Category B. B-59 AC 29.71-1 Autorotational Characteristics – Typical. B-68 AC 29.75-1 Category A Conventional Landing - Clear Heliport. B-78 AC 29.75-2 Category A Vertical Landing. B-79 AC 29.75A-1 Category A Conventional Landing - Clear Heliport. B-89 AC 29.75A-2 Category A Vertical Landing. B-90 AC 29.79-1 Height-Velocity (HV) Diagram. B-98 AC 29.79-2 Altitude/Weight Accountability. B-99 AC 29.175-1 Static Longitudinal Stability. B-115 AC 29.175A-1 Static Longitudinal Stability. B-116.3 AC 29.251-1 Demonstration Points. B-127 AC 29.337-1 Load Factor Gross Weight Curve. C-7 AC 29.351-1 Sample Yaw/Forward Speed Diagram. C-12 Page xix

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AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.351A-1 Sample Yaw/Forward Speed Diagram. C-14 AC 29.351B-1 Yaw/Forward Speed Diagram. C-17 AC 29.351B-2 Safety Factors for Probability of Failure. C-18 AC 29.427-1 (View Looking Forward). C-31 AC 29.427-2 (View Looking Forward). C-32 at Floor Level.

Anthropomorphic Dummy.

Bottom Cushion.

in this guidance.

AC 29.562-1 Seat Restraint System Dynamic Tests. C-58 AC 29.562-2 Schematic Floor Deformation Fixture; Seat Legs Attached C-65 AC 29.562-3 Schematic Test Fixture; Side Wall Mounted Seat. C-69 AC 29.562-4 Installation of Pelvic Lumbar Spine Load Cell in Part 572B C-73 AC 29.562-5 Measurement to be Taken Over Full Width of Seat C-85 AC 29.562-6 Post Test Condition. C-85 AC 29.562-7 Impulse Shape. C-91 AC 29.562-8 Change in Velocity. C-91 AC 29.571-1 Comparison Percent of Time in Level Flight. C-96 AC 29.571B-1 Seven Fatigue Evaluation Methods discussed C-103 AC 29.571B-2 No Growth. C-118 AC 29.571B-3 Slow Growth of Undetectable Crack. C-118 AC 29.571B-4 Slow Growth of Detectable Crack. C-119 AC 29.571B-5 Decreasing Residual Life in Secondary Load Path C-122 for Multiple Element Crack Growth with Inspections.

AC 29.571B-6 Multiple Load Path Structure Analytical Evaluation To C-123 Support Inspection for a Failed Load Path.

AC 29.571B-7 Multiple Load Path Structure Evaluation by Test to C-125 Support Inspection for a Failed Load Path.

AC 29.571B-8 Multiple Load Path Structure Analytical Evaluation to C-126 Support Inspection for Less than a Failed Load Path.

Page xx

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AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.573-1 Schematic Diagram of Building Block Tests. C-144 AC 29.573-2 Characterization of Impact Damage. C-150 AC 29.573-3 Residual Strength vs. Time. C-154 AC 29.573-4 Illustration of Residual Strength and Damage Size. C-155 Relationships for Fail-Safe Substantiation.

AC 29.621-1 Example of Casting Load Sheet Retract Actuator D-24 Support-Landing Gear.

Certifiable Under § 29.865.

Combination vs. Applicable Cargo Type Data And Definition Summary.

AC 29.903C -4 Blade Fragment Definition.

AC 29.621-2 Castings to be Static Tested for Model XYZ. D-26 AC 29.773-1 Cockpit Visibility. D-73 AC 29.801-1 Sea State Code - World Meteorological Organization. D-96 AC 29.863-1 Flammable Fluids Fire Protection Flow Chart. D-128 AC 29.865-1 Summary of Part 133 Rotorcraft Load Combinations D-136 AC 29.865B-1 U.S. Operational (Part 133) Rotorcraft-Load D-146 AC 29.877-1 Continuous Icing-Temperature vs. Altitude Limits. D-181 AC 29.877-2 Intermittent Icing-Temperature vs. Altitude Limits. D-182 AC 29.877-3 Mean Effective Drop Diameter-Microns. D-183 AC 29.877-4 Intermittent Icing-Liquid Water Content vs. Drop Diameter. D-184 AC 29.877-5 Supercooled Clouds. D-185 AC 29.903C-1 Estimated Path of Fragments. E-23 AC 29.903C -2 Single One-Third Disc Fragment. E-24 AC 29.903C -3 Intermediate and Small Pieces of Debris. E-25 E-26 AC 29.903C -5 Distribution of Translational and Rotational Kinetic. E-27 Energy of Rotor-Component Fragments as a Function of Fragment Size.

AC 29.903C -6 Cross Section through Aircraft at Plane of Rotation of E-28 the Engine Disk Fragment.

AC 29.923A-1 Rotor Drive System Endurance. E-48 Page xxi

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AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.923A-2 Takeoff Run Portion of Rotor Drive System Endurance … E-49

Test If 2 ½ Minute OEI Rating is Requested.

AC 29.923A-3 Takeoff Run Portion of Rotor Drive System Endurance … E-50

Test If 30 Second/2 Minute OEI Rating is Requested.

AC 29.931-1 Cantilevered Shaft, First Critical Speed … E-62

AC 29.931-2 Shaft Between Support Bearings, First Critical Speed … E-62

AC 29.952-1 Static Tension and Shear Loading Modes … E-87

AC 29.952-2 Hose Assembly Tests … E-88

AC 29.952-3 Minimum Average and Individual Loads for Hose and … E-89

Hose-End Fitting Combinations.

AC 29.952-4 Typical Method of Breakaway Fuel Fitting Load Calculations. E-90

AC 29.952-5 Typical Methods of Frangible or Deformable Attachment … E-91

Load Calculations (Frangible Bolts).

AC 29.952-6 Typical Methods of Frangible or Deformable Attachment … E-92

Load Calculations (Frangible Baffle).

AC 29.955A-1 Fuel Flow. E-102

AC 29.1041-1 Additional Cooling Test Point … E-139

AC 29.1045-1 All-Engine-Critical Altitude Additional Cooling Test Point … E-146

AC 29.1191-1 Table of Materials and Gages Acceptable for Fireproof … E-190

Protective Devices with Flat Surface Geometries.

AC 29.1309-1 Relationship Between Probability and Severity of Failure.......F-1 7 / F-40.5 Condition Effects.

AC 29.1309-2 Failure Condition Categories and Probability Definitions. .......F-18 / F-40.6

AC 29.1309-3 Failure Condition Conversion … F-19

AC 29.1309-4 Safety Assessment Process Overview....................................F-24 / F-40.10 AC 29.1309-5 Depth of Analysis Flowchart....................................................F-2 5 / F-40.11

AC 29.1321-1 Primary Fields of View … F-43

AC 29.1321-1 Secondary Fields of View … F-44

AC 29.1329-1 Height Loss Method … F-54

AC 29.1329-2 Deviation Profile … F-58

AC 29.1329-3 Deviation Profile … F-60

AC 29.1419-1 Continuous Icing-Temperature vs. Altitude Limits … F-110

Page xxii

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AC 29-2C, Chg 7 2/4/2016 Figure No. Title Page No.

AC 29.1419-2 Intermittent Icing-Temperature vs. Altitude Limits … F-111

Dispensing Equipment Design.

Fail-Safe Substantiation.

AC 29 MG 11-1 Crack Growth for Single Element Structure ............................MG 11-14 AC 29 MG 11-2 Crack Growth for Remaining Elements of Multiple..................MG 11-16 Element Structure.

AC 29 MG 11-3 No Crack Growth Scenario. ....................................................MG 11-18 AC 29 MG 11-4 Slow Crack Growth Scenario. .................................................MG 11-18 AC 29 MG 11-5 Detectable Crack Growth Scenario.........................................MG 11-19 Reduction of Software Level.

with a Ground Base that uses COTS Software for Operational Software

AC 29.1419-3 Continuous Icing-Liquid Water Content vs Drop Diameter … F-112

AC 29.1419-4 Intermittent Icing-Liquid Water Content vs Drop Diameter … F-113

AC 29.MG 3-1 Category II Approach Evaluation.............................................MG 3-5 AC 29.MG 3-2 Category II Approach Evaluation. (cont.) ................................MG 3-6 AC 29.MG 5-1 Acceptable Ultimate Load Factor for Agricultural ....................MG 5-3 AC 29.MG 5-2 Sketch of Tank Pressure Test.................................................MG 5-7 AC 29.MG 6-1 Typical Liquid Oxygen System................................................MG 6-22 AC 29.MG 6-2 Typical Liquid Oxygen System - Using Combination Valve.....MG 6-23 AC 29 MG 8-1 Schematic Diagram of Building Block Tests............................MG 8-16 AC 29 MG 8-2 Characterization of Impact Damage........................................MG 8-20 AC 29 MG 8-3 Residual Strength vs. Time. ....................................................MG 8-23 AC 29 MG 8-4 Residual Strength and Damage Size Relationships for...........MG 8-24 AC 29 MG 9-1 Installed Engine Fuel Flow Capacity Check ............................MG 9-8 AC 29 MG 14-1 Refrigeration Cycle..................................................................MG 14-7 AC 29 MG 15-1 Flow Chart for Use of Mitigating Actions Applied for ...............MG 15-8 AC 29 MG 15-2 Flow Chart for Application of HUMS Specific Software...........MG 15-11 AC 29 MG 17-1 Verification Methodology.........................................................MG 17-4 AC 29 MG 19-1 Primary Field of View ..............................................................MG 19-15 AC 29 MG 19-2 Secondary Field of View .........................................................MG 29-16 AC 29 MG 20-1 Primary and Secondary Horizontal Fields of View ..................MG 20-18 Page xxiii

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AC 29-2C, Chg 7 2/4/2016 AC 29 MG 20-2 Primary and Secondary Vertical Field of View ........................MG 20-19 AC 29 MG 20-3 Sample HF Compliance Matrix................................................MG 20-27 AC 29 MG 21-1 Generic Examples by ATA ......................................................MG 21-5 AC 29.APX B-1 Rotorcraft Dynamic Stability Requirements for IFR............... Apdx B-27 Page xxiv

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Acronyms AC advisory circular ACO aircraft certification office ADF automatic direction finding ADI attitude direction indicator AEH airborne electronic hardware AEO all engines operating AdFC advanced flight control AFCS automatic flight control systems AFGCS automatic flight guidance and control systems AGL above ground level AHRS attitude heading reference system Amdt. Amendment APU auxiliary power unit ATC air traffic control BIM blade inspection method CAM cockpit area microphone CAR Civil Air Regulations CAS calibrated air speed CBIM cockpit blade inspection method CDP critical decision point CFR Code of Federal Regulations CG center of gravity COTS commercial off-the-shelf CPS cycles per second CRFS crash resistant fuel system CRI certification review item CRT cathode ray tube CVR cockpit voice recorder DER designated engineering representative DME distance measuring equipment EASA European Aviation Safety Agency ECAS engine caution advisory systems ECU environmental control unit EDS 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 EVS FAA FADEC FAR FEM FHA FMEA FPM FPS FTA GPS GW HEC HF HIRF HRD HUMS HV IAS ICA ICAO ICS IFR Civil Aviation Equipment enhanced vision system Federal Aviation Administration full authority digital engine control Federal Aviation Regulations [OBSOLETE term.] finite element modeling functional hazard assessment failure mode and effects analysis feet per minute feet per second fault tree analysis global positioning system gross weight human external cargo high frequency high intensity radiated fields high rate of discharge health and usage monitoring systems height-velocity indicated airspeed instructions for continued airworthiness International Civil Aviation Organization inter-communication system instrument flight rules Page xxv

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AC 29-2C, Chg 7 2/4/2016 PCDS personnel carrying device IGE in ground effect Systems IIDS integrated instrument display system ILS instrument landing system IMC instrument meteorological conditions PCF PIO PSA PSIG post crash fire pilot induced oscillation preliminary safety assessment pounds per square inch INS inertial navigation system gauge ISA international standard QPL qualified parts list atmosphere RFM rotorcraft flight manual ISIS integral spar inspection system ITT inter-turbine temperature RFMS rotorcraft flight manual supplement KCAS knots calibrated airspeed RGL Regulatory and Guidance KIAS knots indicated air speed Library KTAS knots true airspeed LDP landing decision point RPM RTCA revolutions per minute Radio Technical Commission LPV localizer performance with Of Aeronautics 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 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 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 PBA pitch bias actuator conditions Page xxvi

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**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 assure coordination with the affected agency and industry elements including the Assistant Chief Counsel. All comments will be considered and the disposition will be documented by the Rotorcraft Standards Staff. ASW -100 will issue the special conditions.

b. Basis for Development.

(1) Special conditions are justified on the basis of the existing Part 29 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 assure they are inadequate or inappropriate in light of a new and novel design feature.

Page FAR 21 - 1

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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 29.

AC 21.33 § 21.33 INSPECTION AND TESTS.

a. Applicant Responsibility. Section 21.33 requires the applicant to:

(1) Assure 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 assure 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 29.

(2) As changes to the rotorcraft are made during the test program, the flight test crew should assure 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 43

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 pract ices that Page FAR 21 - 3

### Page 44

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 45

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

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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 assure 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) Normally the frequency of instrument calibration should not exceed 90 days.

However, 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 one year is typical for strain gauged 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

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**CHAPTER 2. PART 29**

**AIRWORTHINESS STANDARDS**

**TRANSPORT CATEGORY ROTORCRAFT**

SUBPART A - GENERAL AC 29.1. § 29.1 (Amendment 29-21) 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 so that Part 29, in effect, deals with rotorcraft which have a maximum weight greater than 6,000 pounds. In Part 29, there are two categories of rotorcraft, Category A and Category B.

(1) Category A. Category A provides the most rigid rules, requiring multiengine design with independent engines, fuel systems, and electrical systems. Category A design requires that no single failure can cause loss of more than one engine. Although there is no limit on maximum weight, Category A rotorcraft are certificated at a weight which will assure a minimum climb capability in the event of engine failure and with adequate surface area to assure a safe landing in the event an engine fails early in the takeoff run.

(2) Category B. Category B rotorcraft may be single or multiengine and may not have a maximum weight greater than 20,000 pounds. Category B rotorcraft are not required to have the capability for continued flight with an engine failed.

(i) Without Engine Isolation. For single engine rotorcraft and multiengine rotorcraft without engine isolation, the height-velocity diagram is conducted with sudden failure of all engines and the takeoff distance is measured through the clear area of the diagram to the 50- foot point with all engines operating. The landing distance is determined with all engines inoperative.

(ii) With Engine Isolation. Category B multiengine rotorcraft may be certificated with the Category A design features of Part 29. These rotorcraft meet the design requirements of Category A, but the performance requirements of Category B.

Stay-up ability after an engine failure is not assured. The takeoff is conducted with all engines operating, while the height velocity diagram and landing distances are determined with the most critical engine inoperative.

(3) Dual Certification, Categories A and B. A multiengine rotorcraft may be certificated under both categories provided requirements for both categories are met.

This combination will typically result in conditions (1) and (2)(ii) above with the primary differences being the gross weight allowed and the surface areas required for takeoff.

b. Procedures. None.

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AC 29.1A. § 29.1 (Amendment 29-39) APPLICABILITY.

a. Explanation. Amendment 29- 39 revised the reference in § 29.1(e) from §§ 29.79 to 29.87, which is a redesignation of the section number for the height-velocity envelope. 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 so that Part 29, in effect, deals with rotorcraft which have a maximum weight greater than 6,000 pounds. In Part 29, there are two categories of rotorcraft. Category A and Category B.

(1) Category A. Category A provides the most rigid rules, requiring multiengine design with independent engines, fuel systems, and electrical systems. Category A design requires that no single failure can cause loss of more than one engine. Although there is no limit on maximum weight, Category A rotorcraft are certificated at a weight which will assure a minimum climb capability in the event of engine failure and with adequate surface area to assure a safe landing in the event an engine fails anywhere in the flight envelope, including takeoff or landing operations.

(2) Category B. Category B rotorcraft may be single or multiengine and may not have a maximum weight greater than 20,000 pounds. Category B rotorcraft are not required to have the capability for continued flight with one engine inoperative.

(i) Without Engine Isolation. For single engine rotorcraft and multiengine rotorcraft without engine isolation, the height-velocity diagram is conducted with sudden failure of all engines and the takeoff and landing distances are measured with all engines operating.

(ii) With Engine Isolation. Category B multiengine rotorcraft may be certificated with the Category A design features of Part 29. These rotorcraft meet the design requirements of Category A but the performance requirements of Category B.

Stay-up ability after an engine failure is not assured. The takeoff distance is determined with all engines operating. The landing distance, at the option of the applicant, may be determined with the critical engine inoperative or with all engines operating. The height-velocity diagram is determined following failure of the most critical engine.

(3) Dual Certification, Categories A and B. A multiengine rotorcraft may be certificated under both categories provided requirements for both categories are met.

This combination will typically result in conditions (1) and (2)(ii) above with the primary differences being the gross weight allowed and the surface areas required for takeoff.

b. Procedures. The guidance material in paragraph AC 29.1 does not apply to rotorcraft certified with Amendment 29-39 or later.

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AC 29.2. § 29.2 (Amendment 29-32) SPECIAL RETROACTIVE REQUIREMENTS.

a. Explanation.

(1) Amendment 29-32 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 29 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 29-29.

(3) Combined safety belt and harness strength standards system follows:

(i) Those rotorcraft type designs certificated to static strength standards alone prior to Amendment 29-29, 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 29 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 29-29 should use torso restraint systems approved under TSO C114 or approved under equivalent standards such as those contained in Part 29.

(4) Load Distribution and Design Requirements. Although not stated in § 29.2, a 60 percent and 40 percent load distribution between the safety belt and harness, respectively, is required in § 29.785(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 § 29.785(f)(2) of Amendment 29-24 for those rotorcraft with that certification criteria or should include the 1.15 factor as described in § 29.625 (and predecessor § 7.355(c)(2) CAR Part 7) standards for those rotorcraft with the earlier certification criteria. A factor is used whether test results or analysis methods are used for static substantiation of the seating systems. Refer to paragraph AC 29.785b(1)(i) (§ 29.785).

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AC 29-2C, Chg 1 2/12/03

(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 § 29.2 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 § 29.562, 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 tests.

(i) Paragraph 27.562 of this AC concerns in part the dynamic test standards of Amendment 29-29.

(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 § 29.561(b)(3)(iii) is 2.0g, and for later designs (Amendment 29-29 and later), it is 8.0g.

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**CHAPTER 2. PART 29**

**AIRWORTHINESS STANDARDS**

**TRANSPORT CATEGORY ROTORCRAFT**

**SUBPART B - FLIGHT**

GENERAL AC 29.21. § 29.21 (Amendment 29-24) 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 must be shown for each combination 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/AUTHORITY 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 Category A, are conducted to meet specific requirements. Detailed performance test procedures and allowable extrapolation or simulation limits are contained in the respective paragraphs in this order.

(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

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AC 29-2C, Chg 4 9/17/2009 out-of-ground effect (OGE) conditions. From these data the hover ceiling may be calculated.

(ii) Takeoff and landing tests are conducted to determine the total distance to takeoff and land at various combinations of altitude, temperature, and gross weight.

(iii) 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.

(iv) Height-velocity tests are conducted to determine the boundaries of the height versus airspeed envelope within which a safe landing can be accomplished following an engine failure.

(v) Power available tests are conducted to verify or reestablish 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 29.21(a) allows calculation and inference which includes extrapolation and simulation, whereas § 29.21(b) requires demonstration of controllability, stability, and trim. Combinations of §§ 29.21(a) and 29.21(b) may be used to show compliance to 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 insure 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 a(3) above 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 revi2ewed, 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, RPM and cover the range of altitude and temperature for which certification is requested.

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AC 29.21A. § 29.21(Amendment 29-39) PROOF OF COMPLIANCE.

a. Explanation. Amendment 39 added § 29.83 which changes the requirements for determination of landing distance for Category B rotorcraft. This amendment requires landing distance to be determined with all engines operating within approved limits.

b. Procedures. The guidance material presented in paragraph AC 29.21 continues to apply.

AC 29.25. § 29.25 (Amendment 29-12) 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:

Maximum: Structural limits, performance requirements, stability, and controllability requirements.

Minimum: Autorotative rotor RPM, stability, and controllability requirements.

(3) Jettisonable External Cargo.

(i) Paragraph (c) was added by Amendment 29-12 to provide, in the certification standards, a basis for approving an increase in gross weight (exceed standard limits) that would be an external jettisonable load. The attachment device standards were moved from Part 133 (Amendment 133-5) to Parts 27 and 29.

Section 29.865, “External load attaching means,“ now contains the standards, including design features, for the attaching devices. Cargo hoists and hooks were envisioned.

Prior to these amendments, type design approvals were made under Part 133 and the policy in Review Cases Nos. 37 and 55 of FAA Order 8110.6 whenever the standard limits were exceeded.

(ii) In the preamble of Amendment 29-12 (Proposal 2-99, 41 FR 55454, December 20, 1976) the agency stated, in part, that “...§ 29.25(c) is intended to provide only a total weight standard for approving the rotorcraft structure (and propulsion systems) for operation under Part 133.” As indicated in § 29.865, fatigue substantiation of the external cargo attaching means is not required. The rotorcraft structure, rotors, transmissions, engines, etc., are subject to evaluation under § 29.571 for external cargo approval whenever the “standard” structural limitations are exceeded (Review Case Nos. 37 and 55).

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AC 29-2C, Chg 4 9/17/2009

(iii) Whether or not the standard limitations are exceeded, the flight characteristics evaluations/standards of § 133.41 are appropriate even for engineering approval. This Part 133 standard is also applicable for the individual operator to obtain his operating certificate. The operator may use an FAA/AUTHORITY approved RFM supplement for external load operations to prepare a 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 be assured 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 may be warranted. Howev er, if critical areas at minimum flying weights are apparent, extrapolation should not be permitted.

(2) Whenever a gross weight increase (§ 29.25(c)) is requested, a TIA evaluation is necessary to evaluate the new limitations and ensure that § 133.41 for typical or representative cargo shapes and weights (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 the approved limitations are observed. Paragraph AC 29.1581 concerns the RFM and its contents.

(3) See paragraph AC 29.571, § 29.571, for fatigue substantiation and external cargo considerations.

(4) Refer to AC 133-1A, Rotorcraft External-Load Operations in Accordance with FAR Part 133, October 16, 1979, for further information on airworthiness and flight manual policy.

AC 29.25A. § 29.25 (Amendment 29-51) WEIGHT LIMITS.

a. Explanation.

(1) Amendment 29-51 added a new paragraph (a)(4) that requires that the operating envelope for the controllability demonstrated under § 29.143(c) be included in the limitations section of the Rotorcraft Flight Manual (RFM). The change allows, in addition to the 17-knot controllability requirements, the applicant to provide additional Page B - 4

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controllability information within an applicant selected limited azimuth range if the rotorcraft is certified with nine or less passenger seats. This effectively allows increased weights within this limited range. Amendments 29- 21 and 29-24 allowed for this relief and subsequent regulatory policy recognized these limitations as they are now required.

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.

(2) The explanation regarding the relief for presentation of hover controllability limits in AC 29.143.a.(2)(ii) (Amendment 29-24) is superseded by this change.

b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect.

AC 29.27. § 29.27 (Amendment 29-3) 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 of 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 lateral Page B – 5

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AC 29-2C, Chg 4 5/1/2014 loadings resulting in the necessity to establish lateral center of gravity limits. There are two characteristics which may be seriously affected by loading outside the established center of gravity limits; these are stability and control. The established center of gravity limits must be such 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 § 29.143(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. Additional items which may influence the maximum aft center of gravity limits may be malfunctions of automatic stabilization equipment, excessive rotorcraft attitudes during critical phases of flight, or adequate control power to compensate for an engine failure.

(3) Lateral center of gravity limits have become more critical because of the ever increasing utilization of the rotorcraft for such things as unusual and unsymmetric lateral loads, both internal and external. Maximum allowable lateral center of gravity limits have also influenced the results of the unusable fuel determination.

(4) Summarizing, 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.

AC 29.29. § 29.29 (Amendment 29-15) 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 non- permanently 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 certificated 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 § 29.29 is accomplished by the use of an Page B – 6

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equipment list specifying the installed equipment at the time of weighing and the weight moment arm 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 procedure 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 the current version of 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 or rotorcraft flight manual supplement.

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 that 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.

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AC 29-2C, Chg 4 5/1/2014 AC 29.31. § 29.31 REMOVABLE BALLAST.

a. Explanation. This regulation provides the option of using removable ballast for operational flights to obtain center of gravity locations that are in 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 permit its use on transport rotorcraft. If removable ballast is used, the rotorcraft flight manual must include instructions regarding its use and limitations.

See paragraph AC 29.873 for information on ballast provisions.

b. Procedures. None.

AC 29.33. § 29.33 (Amendment 29-15) MAIN ROTOR SPEED AND PITCH LIMITS.

a. Explanation.

(1) General. This rule 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.

A power-on range of approximately 3 percent has in the past been the minimum range 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, but one fixed value may suffice. Transient power -on values may also be acceptable provided they are substantiated.

(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 operation environment; i.e., at altitudes between approximately sea level and 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 extreme cold temperatures, the normal low pitch setting may not be sufficient to maintain autorotational rotor speed values within Page B – 8

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limits. If this oc curs, 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 down rigging 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) Additional RPM Ranges. Some applicants have elected to certify their aircraft with additional RPM ranges in an attempt to realize additional performance during certain flight conditions or maneuvers such as Category A OEI continued and rejected takeoffs and balked landings. Such additional RPM ranges have been found acceptable as long as all pertinent FAR requirements are fully substantiated for operation in that range. The substantiation should include drive system endurance and flight test verification of performance and flight characteristics during applicable maneuvers, in the additional RPM range. The FAA/AUTHORITY does not define additional RPM ranges as transient since all applicable requirements must be satisfied for approval of that range.

(5) Low Speed Warning. If it is possible under expected operating conditions for the rotor speed to fall below the minimum approved values, the requirement exists for a low rotor speed warning. This warning is required on all single- engine rotorcraft and on multiengine rotorcraft where there is not an automatic increase in remaining engine(s) power output upon failure of an engine. Although today’s multiengine rotorcraft do not require a low rotor speed warning according to the rule, essentially all have warning systems 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 rotorcraft has a warning system cutout if the collective is full down, and others 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 29.1509 (§ 29.1509) for additional information on rotor limits determination and testing.

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AC 29-2C, Chg 4 5/1/2014

**SUBPART B - FLIGHT**

PERFORMANCE AC 29.45. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation.

(1) Changes to various part 29 sections, which did not include amending § 29.45, added new and revised airworthiness standards for the performance of transport category rotorcraft and renumbered several sections within the performance section of Subpart B. The performance section of this guidance material has been organized for easy use with rotorcraft certificated before or after this amendment. To achieve this, some of the guidance material has been duplicated under different paragraph numbers.

A statement at the beginning of each of these paragraphs indicates where other pertinent information can be found.

(2) Section 29.45 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 applicable test conditions. The test results should be analyzed and expanded by an approved methodology. The guidance within this section is considered an approved methodology.

(3) Performance should be based on approved engine power as determined in paragraph b.(4) below and not on any transient limits. Approved transient limits are basically for inadvertent overshoots of approved operational limits. Any sustained operation in these transient limit areas usually require some form of special maintenance. However, for such demonstrations as rejected and continued OEI category A takeoffs and HV determination, low rotor speeds have been authorized based upon additional structural and drive system substantiation (see section 29.33 of this AC).

(4) 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.

(5) As defined in 14 CFR § 1.1, the 30-second and 2-minute OEI power ratings are based on up to three periods of use during a single flight. The purpose of the three applications 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 Page B – 10

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the use of these time-limited power ratings is only permitted once in each of the above three uses (i.e., 30-second power must not be used more than once during the initial recovery from an engine failure).

(6) 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.

(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 may be tolerated for takeoff and landing distance tests but not for hover performance.

Likewise higher winds may be tolerated during hover performance testing on large, heavy rotorcraft with high rotor downwash velocities than for smaller rotorcraft with lower 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) Past experience has shown that a steady wind of 0 to 10 knots will result in acceptable takeoff and landing performance if distances are corrected for the winds measured during these tests. This is not the case for vertical takeoffs and landings. To obtain consistent and repeatable vertical performance data, using wind speeds up to 5 knots is acceptable. In actuality, a rotorcraft may exhibit reduced IGE hover performance in wind speeds from 3 to 15 knots due to partial immersion of the main rotor in its own vortex. Since the height -speed envelope determination is affected by wind just as vertical takeoff and landing performance are, the same allowable winds for testing should be adhered to for HV testing; i.e., 0 to 5 knots. For category A testing, the effects of crosswind and tailwind should also be considered up to the maximum for which category A certification is requested.

(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) Conventional takeoff and landing - 0 to 10 (data to be corrected) (C) Vertical takeoff and landing - 0 to 5 knots (D) HV - 0 to 5 knots Page B – 11

### Page 62

AC 29-2C, Chg 4 5/1/2014

(iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests. The wind effects on required runway length for takeoff and landing distances may be shown in the flight manual.

(v) Full wind credit may be given for conventional takeoff and landing field lengths. This credit should not be more than the nominal wind component along the takeoff or landing path opposite to the direction of flight.

(2) Altitude Effects: Extrapolation and Interpolation.

(i) Using FAA/AUTHORITY approved methodology:

(A) Hover, takeoff, and landing performance may be extrapolated from test data up to a maximum of ±4,000 feet density altitude from the test altitude.

(B) Experience has shown that 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 of 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.

(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.

(i) 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 § 29.1527 and delineates the maximum altitude up to which operation is allowed. This altitude normally constitutes the maximum cruise or enroute altitude.

(B) Maximum weight, altitude, and temperature for takeoff and landing constitute a limitation. The maximum takeoff and landing altitude may be coincident with but never above the maximum operating altitude limitation. Takeoff and landing, hover ceiling data, and presentation requirements are presented in §§ 29.51, 29.53,

29.59, 29.63, 29.73, 29.1583, and 29.1587.

Page B – 12

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(ii) Procedures.

(A) In establishing the maximum takeoff and landing altitude, the following tests are normally required:

(1) Takeoff (§§ 29.51-29.63)

(2) Climb (§§ 29.64-29.67)

(3) Performance at minimum operating speed (§ 29.49)

(4) Landing (§ 29.75)

(5) HV envelope (§ 29.87)

(6) IGE controllability (§ 29.143(c))

(7) Cooling (§§ 29.1041-29.1045)

(8) Engine operating characteristics (§ 29.939) 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.

(C) Prior to Amendment 29-21, HV information was an operating limitation.

With the adoption of Amendment 29-21, the HV curve is performance information for category B rotorcraft with nine or less passenger seats but remains a limitation for category A rotorcraft and category B rotorcraft with 10 or more passenger seats.

(D) Prior to Amendment 29-24, IGE controllability was required in 17 knots of wind to the maximum takeoff and landing conditions. With the adoption of Amendment 29-24, if IGE or OGE hover performance is presented for a category B rotorcraft to an altitude in excess of that for which IGE controllability at 17 knots is presented, the maximum safe wind demonstrated for hover operations must be presented in the RFM. The amendment did not change the requirement for category A rotorcraft.

(E) 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 29.45-1 and AC 29.45-2.

Page B – 13

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AC 29-2C, Chg 4 5/1/2014 (F) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in paragraph b.(2) above and not above the lowest limiting altitude resulting from the requirements listed in paragraph b.(3)(ii)(A) above.

(4) Temperature Effects.

(i) Background.

(A) The regulations prohibit any unsafe design feature throughout the range of environmental conditions for which certification is requested. The regulations also require that the performance and handling qualities be determined over the approved range of atmospheric variables selected by the applicant.

(B) Substantiation of temperature effects on performance and handling characteristics is required throughout the approved temperature range. 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 blade 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.

(C) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected, sometimes significantly, in both the performance and flying qualities areas. Hot temperature conditions although not shown to be as critical should be given consideration.

(D) 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 § 29.1309(a), the applicant needs to substantiate the total rotorcraft at the extreme temperatures for which certification is requested.

(ii) Procedures.

(A) The FAA/AUTHORITY is responsible for verifying the applicant’s predictions of performance and handling characteristics at the temperature extremes for which certification is requested. A limited flight verification, if necessary, could include spot checks of hover and climb performance, IGE controllability, roughness determination, simulated power failure, static stability, height-velocity, VNE/VD evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.

Page B – 14

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(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) Weight Effects. Test weights should be maintained within +3 percent and -1 percent of the target weight for each data point. Weight may be extrapolated or interpolated only along an established W/σ line within the allowable altitude extrapolated range.

(6) 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 horsepower used to complete the flight manual performance must be based on horsepower 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 horsepower values at specification temperatures and speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specified horsepower 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 horsepower. 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 29.45-3 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 horsepower values 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 Page B – 15

### Page 66

AC 29-2C, Chg 4 5/1/2014 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 29.45-4.

(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 horsepower resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet or exhaust design. The rotorcraft manufacturer usually conducts test to confirm the installed specification. Methods used vary widely between manufacturers, but usually include some combination of ground and flight tests. Figure AC 29.45-5 is a typical example of an installed power available chart for one set of conditions.

(F) This predicted installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion decreases the available horsepower output and 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 horsepower available from a minimum installed specification engine in a hover. Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual horsepower available from a minimum installed specification engine at the published climb airspeed.

This will allow the manufacturer to take advantage of, for example, increased inlet efficiency.

(ii) Procedures.

(A) To this point the minimum installed specification engine horsepower output has been predicted and calculated for various flight conditions. It is imperative that the predicted values be verified by actual flight test. The flight test involves obtaining engine performance measurements at various power settings, altitudes, and ambient temperatures. The data should be obtained at the actual flight condition for which the performance is to be presented (i.e., hover, climb, or cruise).

(B) Following an initial application of power, 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 is held constant. This is a characteristic Page B – 16

### Page 67

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 applicable value (takeoff, 30-second, or 2 1/2-minute power), the engine is first stabilized at a low power setting. After stabilization, rapidly increase the power demand to takeoff, 30-second and 2 1/2-minute power levels as necessary. 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 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 1/2-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. An engine may be certified to produce a specification horsepower at a particular temperature or engine speed rating with higher maximum limit value approved. Only the rating values should be used to determine the installation losses.

The limit values of engine temperature and speed are established and certified to allow specification powers to continue to be developed as the engine deteriorates in service.

(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 installed specification engine installed on a production rotorcraft is capable of producing the horsepower 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 a total of the different minimum installed specification horsepower values. 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, etc., must be accounted for accordingly.

Page B – 17

### Page 68

AC 29-2C, Chg 4 5/1/2014 (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 minimum 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, and 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 29.45-6 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 pre-takeoff 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 horsepower values. This occurs when the curve of measured corrected horsepower 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.

(7) 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 – 18

### 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 depreciated engine power. These include:

(1) Engine characteristics (§ 29.939). The reduced compressor discharge pressure, PC, would reduce engine surge margin and possibly affect engine response and engine air-restart capability. These items 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 producers 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 affect 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.(6) (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.(6) (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 cold and high; bleed valves and other devices which schedule on gas producer Page B – 19

### Page 70

AC 29-2C, Chg 4 5/1/2014 discharge pressure are most likely to function and reduce power on a hot day at low altitude.

(C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification region established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated procedures will be introduced for 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.(6) 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.

(D) If a complex power assurance procedure is presented with involved data reduction and trending requirements, consideration should be given to restricting the use of deteriorated power to operators where close control over operations is exercised or the operator has demonstrated the ability to operate safely with deteriorated engines.

(8) Engine Failure Testing Considerations

(i) For all tests to examine behavior following an engine failure, usually the failure of the engine is simulated in some way. For engines with a hydro-mechanical governing system, it is common practice to close the throttle quickly to idle. For rotorcraft equipped with a FADEC, and particularly those with a 2 minute/30 second OEI rating structure, it is common practice to simulate an OEI condition by using reduced power on all engines by means of a flight test tool.

(ii) In every case, it must be demonstrated that all aspects of rotorcraft and powerplant behavior are identical to those that would occur in the event of an actual engine failure with the remaining engine developing minimum-specification power. Of particular concern are “dead engine” power decay characteristics, “live engine” acceleration characteristics, and rotor RPM control.

(iii) To this end, it is expected that a number of actual engine shut down tests will be conducted to generate sufficient data to validate the fidelity of the flight test tool and methodology, which will then allow its use in developing regulatory performance data. In general, it is best to conduct the tests in a low hover with the rotorcraft stabilized below the HV low point. An engine is then shut down and, following the appropriate pilot intervention time, the collective is raised to cushion the landing.

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AC 29.45A. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation. Amendment 29-24 adds § 29.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered aircraft. 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 29.45B. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation. Although § 29.45 was not changed by Amendment 29-34, 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 material on power assurance procedures to ensure that the OEI power level can be achieved. All of the guidance material pertaining to AC sections 29.45 and 29.45A remain in effect.

AC 29.45C. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.

a. Explanation. Although § 29.45 was not changed by Amendment 29-51, that amendment added new performance and handling qualities requirements for transport category rotorcraft. Included within these regulatory changes is OGE handling qualities.

Additionally, hover performance requirements were re-identified from § 29.73 to § 29.49.

b. Procedures. All of the guidance material pertaining to AC sections 29.45,

29.45A, and 29.45B 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 § 29.73 is applicable to § 29.49.

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AC 29-2C, Chg 4 5/1/2014

CERTIFICATION BASIS 14 CFR pa rt 29 CAR 7 Rqms 29-Amdt. 21 29-Amdt. 1 Original Original HV CAT A Cat A & B (>9 Cat A: Cat A: Cat. A:

Ref. TEST pax seats): 1. MGW Sea 1. MGW Sea 1. MGW Sea

29.25 CONDITIONS 1. W.A.T. for Level Level Level

29.87 which t.o. and 2. Max. IGE 2. Max. IGE 2. Max. IGE wt.

29.1517 ldg. are wt. Max. wt. Max. Max.

29.1581 approved. alt. capability. alt. capability. alt. Capability.

29.1583 2. Failure of 3. Failure of 3. Failure of 3. Failure of

7.11 critical engine. critical engine. critical critical engine.

7.715 engine.

7.741 AC 29- CAT A 3. HV is 4. HV is 4. HV is 4. HV is 2C, RFM limitation. limitation. limitation. limitation.

Sections 4. Type of ldg. 5. Type of ldg. 5. Type of ldg. 5. Type of ldg.

29.45 & surface. Surface. surface. surface.

**29.79 CAT B TEST**

CONDITIONS Cat B (</=9 pax seats):

1. MGW Sea Level

2. Max. OGE wt. Max. alt.

Capability

3. Complete power failure, or failure of critical engine (w/eng isolation).

Cat B:

1. MGW Sea Level

2. Max. IGE wt. Max. alt.

Capability

3. Complete power failure, or failure of critical engine (w/eng isolation).

Cat B:

1. MGW Sea Level

2. Max. IGE wt. Max. alt.

Capability

3. Complete power failure, or failure of critical engine (w/eng isolation.)

Cat B:

1. MGW Sea Level

2. Max. IGE wt.

Max. alt.

Capability

3. Complete power failure, or failure of critical engine (w/eng isolation.)

CAT B 4. HV is perf. 4. HV is 4. HV is 4. HV is RFM Info.

5. Type of ldg.

surface.

limitation info.

5. Type of ldg.

Surface.

limitation info.

5. Type of ldg.

surface.

limitation info.

5. Type of ldg.

surface.

FIGURE AC 29.45-1. HV REQUIREMENTS Page B – 22

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CERTIFICATION BASIS 14 CFR part 29 CAR 7 Requirement s 29-Amdt. 24 29-Amdt. 3 Original Original IGE

CONTROL Ref. 29.25

29.1583

7.121

**7.743 AC 29-2C CAT A TEST**

CONDITION S Cat A

1. W.A.T. for which

t.o. and ldg.

are approved.

2. Critical wt.

Critical CG Critical Nr

3. Wind not less than 17 kts.

1. Conditions selected by the applicant.

2. Critical CG Critical Nr

3. Wind not less than 17 kts.

1. Conditions selected by the applicant.

2. Critical CG Critical Nr

3. Wind not less than 20 mph.

1.

Conditions selected by the applicant.

2. Critical CG Critical NR

3. Wind not less than 20 mph.

CAT A 4. Max. 4. Max safe 4. Max safe 4. Max.

Sections

29.45 and

29.143 RFM allowable wind is limitation.

wind above max. alt. For which 17 kt.

Wind envelope is established is perf. info.

wind above alt.

for which 17 kt.

wind envelope is established is perf. info.

allowable wind above the altitude for which 20 mph wind envelope is est. is perf.

info.

CAT B Cat B:

TEST 1. W.A.T. for CONDITION which t.o. and S ldg. Are approved.

2. Critical wt.

Critical CG Critical Nr

3. Wind speed & quad selected by the applicant.

**CAT B RFM**

4. Max. safe wind is perf.

info.

FIGURE AC 29.45-2 IGE CONTROLLABILITY REQUIREMENTS Page B – 23

### Page 74

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### Page 75

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### Page 77

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V I , ,/ V V / ,v I V V _/ ..

V V I V I I 0 20 40 60 / V / / , / / V / / V / _,, v / /v / / 80 100

**PERCENT TORQUE / ,/ , / / / / 120 10000 FT 8000 FT 60()() FT 4000 FT 2000 FT**

SEA LEVEL 5/1/2014 AC 29-2C, Chg 4 Page B – 26.1

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AC 29-2C, Chg 4 5/1/2014 AC 29.49. § 29.49 (Amendment 29-39) PERFORMANCE AT MINIMUM OPERATING SPEED.

(For performance at minimum operating speed and for hover performance prior to Amendment 29-39, see § 29.73 and section 29.73 of this AC.)

a. Explanation.

(1) Amendment 29-39 re-designated § 29.73 as § 29.49 to relocate the requirements for rotorcraft hover performance. For the purpose of this manual, 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.

(2) Under § 29.49, hover performance should be determined at a height consistent with the takeoff procedure for category A rotorcraft and in ground effect (IGE) for category B rotorcraft. Additionally, out of ground effect (OGE) hover performance should be determined for both category A and B rotorcraft. 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 weights. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle 1 ½ 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 non-dimensional power coefficients (Cp) and thrust coefficients (Ct) for normalizing and presenting test results, a minimum amount of data are required to cover the rotorcraft’s performance 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 landings. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.

(5) The minimum hover height 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 section 29.51 of this AC for the procedure to determine the minimum allowable hover height.

b. Procedures.

(1) Two methods of acquiring hover performance data are the tethered and 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 Page B – 26.2

### Page 79

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 stabilized point. The total thrust the rotor produces is the rotorcraft’s gross weight, 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 cable and load cell are perpendicular to the ground. To insure 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. This can also be accomplished by attaching two accelerometers to the load cell which sense 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. Accurate load cell values may also be obtained by measuring cable angles and, through geometry, determining a corrected load cell value. 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 3 knots or less. Large 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.

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AC 29-2C, Chg 4 5/1/2014

(ix) Hover height.

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 paragraph b.(4) below 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 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) Comprehensive hover performance tests are typically conducted at low, intermediate (approximately 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 section

29.45.b.(2) of this AC 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 approving airworthiness authority a method to provide a reliable hover reference, it is acceptable to conduct OGE tests without ground reference.

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AC 29.51. § 29.51 TAKEOFF DATA - GENERAL.

a. Explanation. Section 29.51 details the conditions under which takeoff performance data can be obtained and presented in the FAA/AUTHORITY approved flight manual. The flight manual must also contain the technique(s) to be used to obtain the published flight manual takeoff performance. Technique should not be confused with exceptional pilot skill and/or alertness as mentioned in § 29.51. Rotorcraft are different from one another and due to this, 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 and used to achieve the performance 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.

b. Background.

(1) Certain special takeoff techniques are necessary when a rotorcraft is unable to takeoff 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, altitude, and temperature for which certification is requested.

(2) The primary factor which determines the rotorcraft’s takeoff performance is the amount of excess power available. Excess power available is the difference between the power required to hover at the reference height above the ground and the takeoff power available from a minimum installed specification engine. Utilizing the total power available to execute a takeoff may not be operationally feasible due to such items as HV constraints. In such situations, hover power required plus some power increment may be the maximum that can be used and the resulting performance determined accordingly.

(3) Landing gear height above the ground should not be greater than that demonstrated satisfactorily for HV, rejected takeoff, and that height for which IGE hover performance data is presented in the RFM, or less than that height below which ground contact may occur when accomplishing takeoff procedures. For rotorcraft fitted with wheels, a running takeoff procedure may be accepted. The hover reference height is established as the minimum landing gear height above the takeoff surface, from which a takeoff can be accomplished consistently in zero wind without contacting the runway.

Category B takeoff must be accomplished with power fixed at the power required to hover at the reference height (not greater than the height for which IGE performance data is presented).

c. Procedure. There are different techniques which may be used in order to 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 Page B – 29

### Page 82

AC 29-2C, Chg 4 5/1/2014 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 schedule airspeed(s) and constant rotor RPM. Power adjustments may be accomplished to maintain targeted power except where procedure requires high workload outside cockpit (i.e., that portion of takeoff where 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).

AC 29.51A. § 29.51 (Amendment 29-39) TAKEOFF DATA - GENERAL.

a. Explanation. Amendment 39 added takeoff requirements in new §§ 29.55, 29.60,

29.61 and 29.62.

b. Procedures. The guidance material presented in paragraph AC 29.51 continues to apply.

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AC 29.53. § 29.53 TAKEOFF: CATEGORY A.

a. Explanation.

(1) A Category A takeoff typically begins with an acceleration and/or climb from a hover to a critical decision point. The rule requires that the critical decision point (CDP) be defined for the pilot in terms of an indicated altitude and airspeed combination. However, other parameters to define the CDP have been accepted by the FAA/AUTHORITY on an equivalent safety basis. A regulatory project has been established to change the rule permitting other parameters to be used for CDP definition.

(2) The requirement to define CDP as a combination of both airspeed and height above the takeoff surface is based on a minimum required total energy concept.

A specific minimum combination of kinetic energy (airspeed) and potential energy (height) must be attained at the CDP to be assured that a continued takeoff can be accomplished following the complete failure of one engine. In § 29.53(b), CDP is required to be “…a combination of height and speed selected by the applicant…” Any other method proposed to define CDP must provide the same level of safety as would be obtained using an airspeed-height combination. When using “time,” “height,” or “airspeed” only as alternative methods of identifying the CDP, they must be combined with a precisely defined takeoff path and crew procedure in order to provide the required equivalent level of safety. In addition, it must be demonstrated that the pilot technique used during the takeoff sequence is easily repeatable and consistently produces the required energy (i.e., airspeed and altitude combination) when the CDP “time,” “height,” or “airspeed” is attained. This condition should be verified during the flight test program.

(3) If an engine fails at the CDP or at any point in the takeoff profile prior to attaining CDP, the rotorcraft must be able to land safely within the established rejected takeoff distance. Flight testing to determine the Category A rejected takeoff distance is very similar to height-velocity testing and should be approached with caution. The initial Category A takeoff profiles should be outside of the Category B height-velocity envelope. Previous programs have shown the low speed point immediately after application of power to be particularly critical.

(4) If an engine fails at the CDP or at any subsequent point in the Category A takeoff profile, a continued safe climb-out capability is assured. The continued takeoff for conventional Category A runway profiles is designed to allow acquisition of the takeoff safety speed (VTOSS), at a minimum of 35 feet above the takeoff surface and a positive rate of climb. During the continued takeoff profile, the pilot is assumed to be flying the rotorcraft via the primary flight controls (cyclic stick, collective, and directional pedals). Manipulation of the throttle controls or beep switches may be permitted as long as such manipulation can be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls. These manipulations of engine controls should not make major adjustments in power, and Page B - 31

### Page 84

should not occur before attaining VTOSS. In no case should this be less than 3 seconds after the critical engine is made inoperative.

(5) Both the rejected takeoff distance and the continued takeoff distance must be determined. Although 29.59(c) suggests a balanced field length requirement, this was not intended. Both rejected and continued takeoff distance should be included in the RFM performance with information stating that the longer distance determines the length of the required takeoff surface. Operations approvals can then determine the required takeoff surface (including stopways and clearways) appropriate for the specific operation.

(6) A typical Category A takeoff profile, assuming an engine failure at the CDP, is shown in figure AC 29.53-1.

b. Procedures.

None.

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100 HEIGHT ABOVE GROUND (FEEl) 35 I I ,I I I VToss I ----- -r- 1 I I :-:-:;:_ STOP ...

A LL ENG INES OPERATING ~TRANSITION SEGMENT

TAKEOFF SEGMENT RTO SEGMENT ----1► 1 ~---DEMONSTRATED SA FE OEI LANDINGS ----1.i

HORIZONTAL DISTANCE .._TAKEOFFCLIIAB S EGMEHT--..;

I FIGURE AC 29.53-1 TAKEOFF PERFORMANCE CATEGORY A 9/30/99 AC 29-2C Page B - 33

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AC 29.53A. § 29.53 (Amendment 29-39) TAKEOFF: CATEGORY A.

a. Explanation. Amendment 29- 39 separated in the text, the Category A takeoff requirement from the definition of a decision point. Category A takeoff performance must be scheduled so that:

(1) If an engine failure is recognized at the Takeoff Decision Point (TDP) or at any point in the takeoff profile prior to attaining TDP, the rotorcraft must be able to land safely within the established rejected takeoff distance. Flight testing to determine the Category A rejected takeoff distance is very similar to height-velocity testing and should be approached with caution. The initial Category A takeoff profiles should be outside of the avoid area of the Category B height-velocity envelope. Previous programs have shown the low speed point immediately after application of power to be particularly critical.

(2) If an engine failure is recognized at the TDP or at any subsequent point in the Category A takeoff profile, a continued safe climb-out capability must be assured.

The continued takeoff for conventional Category A runway profiles is designed to allow acquisition of the takeoff safety speed (VTOSS) at a minimum of 35 feet above the takeoff surface and a positive rate of climb.

(3) Both the rejected takeoff distance and the continued takeoff distance should be determined. A balanced field length is not required by the regulation. Both rejected and continued takeoff distance should be included in the RFM performance section.

Operations approvals can then determine the required takeoff surface (including stopways and clearways) appropriate for the specific operation.

(4) A typical Category A takeoff profile, assuming an engine failure prior to the TDP, is shown in figure AC 29.53A-1.

b. Procedures. None.

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### Page 87

1000 ft.

I ::::i g -cJ C :::, 0 ~ C) Q) > 200 ft 0 .0 ~ ~ -.s:::.

H-V AVOID (29 87) 1/.//1//.((/(,f /

0) ·;;;

I 35 ft 15 ft.

Oft. I+-- Rejected T.O. Dist (29.62) I+-- T.O. Dist (29.61)

v. ?.100 fpm

(29.67 a.1) ,,- Vv .:.150fpm

(29.67a.2) §-'" "',§$ V,oss .._l;- 35 ft.

- ' 15 ft., ' -----.. _____.. Horizontal Distance FIGURE AC 29.53A-1 TAKEOFF PERFORMANCE CATEGORY A 1000 ft 9/30/99 AC 29-2C Page B - 35

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AC 29.55. § 29.55 (Amendment 29-39) TAKEOFF DECISION POINT:

**CATEGORY A.**

a. Explanation.

(1) Amendment 29-39 added a new § 29.55 to redefine the TDP (previously called the CDP) and contained in § 29.53; it further removed the requirement to identify the TDP by height and airspeed, since height alone or other factors may be more appropriate. A Category A takeoff typically begins with an acceleration and/or climb from a hover to TDP. The rule requires that the TDP be defined for the pilot in terms of no more than two parameters such as an indicated height and airspeed combination.

(2) The definition of the TDP is based on a minimum required total energy concept. A specific minimum combination of kinetic energy (airspeed) and potential energy (height) should be attained at the TDP to ensure that a continued takeoff can be accomplished following the complete failure of one engine. In § 29.55(b), TDP is required to be defined by no more than two parameters. When using a single parameter such as time, height, or airspeed as a method of identifying the TDP, the identification must be combined with a precisely defined takeoff path and crew procedure to provide the required equivalent level of safety. In addition, it should be demonstrated that the pilot technique used during the takeoff sequence is easily repeatable and consistently produces the required energy (i.e., airspeed and height combination) when the TDP time, height, or airspeed is attained. This condition should be verified during the flight test program.

b. Procedures. None.

AC 29.59. § 29.59 (Amendment 29-24) TAKEOFF PATH: CATEGORY A.

a. Explanation. The Category A concept limits the rotorcraft takeoff weight such that if an engine failure occurs at or before the CDP, a safe landing can be made or if the engine fails at or after the CDP, the takeoff can be continued. The purpose of these tests is to define the CDP, evaluate the necessary pilot techniques, and determine the required takeoff area for either alternative. The condition of equal distances for either stopping or continuing the takeoff is called a “balanced” field length. The combination of altitude and speed at the CDP which produces a balanced field length is not required for certification. This section deals with the Category A takeoff and rejected takeoff profiles. The profiles necessarily involve consideration of an average pilot skill level as well as a sequence in which it is assumed various configuration adjustments are made to the rotorcraft.

(1) Takeoff. The Category A takeoff path begins with an all -engines-operating acceleration segment to the CDP and continues with a one-engine-inoperative acceleration to takeoff safety speed (VTOSS). (See Conventional Takeoff Profile, figure AC 29.53-1, paragraph AC 29.53.) CDP is a “go/no -go” condition which is analogous to V1 speed in transport airplanes. Prior to CDP the pilot is “stop” oriented, Page B - 36

### Page 89

and when an engine fails in this portion of the takeoff, he will abort because he has not yet achieved sufficient energy to assure continued flight. At the CDP the pilot becomes “go” oriented and when an engine fails at or beyond this point he will continue the takeoff because he no longer has sufficient surface area to abort the takeoff. The takeoff flight path and the CDP must be defined such that a safe landing can be made from any point up to the CDP. This profile may differ significantly from the takeoff flight path developed for Category B weights. The CDP is the last point in the takeoff profile at which a rejected takeoff capability within the scheduled takeoff surface distance is assured. If an engine failure does not occur, the pilot continues the climb and accelerates past the CDP to the recommended climb speed.

(2) Rejected Takeoff. The rejected takeoff profile begins with an all engine acceleration segment to the CDP and ends when the rotorcraft is brought to a complete stop on the designated takeoff surface. The critical engine is made inoperative at the CDP and the landing must be made with the remaining engine(s) operating within approved limits. The rejected takeoff distance is normally measured at a given reference point on the rotorcraft from the start of the takeoff to the same reference point after the rotorcraft has come to a complete stop. This distance should be increased by the rotorcraft length (including main and tail rotor tip paths).

(3) Takeoff Climbout Path.

(i) The “OEI transition segment” is defined as the segment from CDP where the engine becomes inoperative to VTOSS. It is assumed that the maximum approved OEI power is used until the allowable time duration for that power is exhausted. It must be possible for the crew to fly the rotorcraft to V TOSS and attain an altitude of 35 feet and then climb to 100 feet above the takeoff surface by flying the rotorcraft solely by the primary flight controls (including collective). The landing gear may be retracted after attaining a height of 35 feet above the takeoff surface, a speed of VTOSS, and a positive rate of climb. Flight manual procedures may recommend adjustment of auxiliary controls to improve OEI performance. However, compliance with the performance requirements of § 29.67(a)(1) should not be based on use of secondary engine controls such as beepers, etc. Manipulation of the throttle controls or beep switches may be permitted for compliance with the performance requirements of § 29.67(a)(2) as long as such manipulation can be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls. These manipulations of secondary engine controls should not make major adjustments in the power, and should not occur before attaining VTOSS. There should be a minimum delay of 3 seconds after the critical engine is made inoperative before adjustment of secondary engine controls is allowed during the takeoff path determination. The failure of one engine cannot affect continued safe operation of the remaining engines or require any immediate action by the crew per § 29.903(b). If a 2 ½-minute power rating is used, it should be possible to complete the Category A takeoff profile (assuming an engine failure at CDP), accelerate to VTOSS, attain 35 feet above the surface, and complete landing gear retraction prior to exhausting the 2 ½-minute time limit.

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(ii) The takeoff safety speed, VTOSS, is a speed at which 100 FPM rate of climb is assured under conditions defined in § 29.67(a)(1). The takeoff distance is the distance from initial hover to the point at which VTOSS and 35 feet in a climbing posture are attained.

(4) Continued Climbout Path. Continued acceleration and climb capability from 100 feet above the takeoff surface is assured by the 100 FPM VTOSS climb requirement of § 29.67(a)(1) and the 150 FPM requirement of § 29.67(a)(2), normally demonstrated at VY . It should be shown that the rotorcraft can be accelerated from V TOSS to VY in a continuous maneuver without losing altitude, including any configurative change (landing gear retraction, etc.).

b. Procedures.

(1) Instrumentation. A photo theodolite, grid camera, or other position measuring equipment is required together with a ground station to measure wind, OAT, humidity (if applicable), and a two-way communication system to coordinate activities with the aircraft. A crash recovery team with support of a fire engine is highly desirable.

Aircraft instrumentation should record with a time scale: engine parameters (speed, temperature, and power), rotor speed, flight parameters (airspeed, altitude, and normal acceleration as a minimum), flight control positions, power lever position, and landing gear loads. Additionally, a method should be devised to allow correlation of the aircraft instrumentation data with the space position data to accurately determine the length of the various takeoff segments.

(2) Establishing the Critical Decision Point (CDP).

(i) The CDP should be definable with the minimum crew using standard cockpit instrumentation. If a radar altimeter is used, it should be included in the minimum equipment list. If barometric altitude is used to define CDP, the operating conditions at which the altimeter is set should be defined. This is normally done on the ground with the minimum collective pitch. If the wind influences the altimeter reading, the correct relative wind information should be provided. Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the CDP becomes largely a function of the surface area required for takeoff. If takeoff conditions scheduled include considerable surface area (on the order of 2,000 feet), the CDP airspeed may be a high value near VY. This will allow a higher takeoff weight and demonstrate compliance with the VTOSS climb requirement of § 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If required surface area is a small value, CDP will necessarily be some lower airspeed value to allow for an aborted takeoff on the available surface. Weight may need to be reduced at lower values of CDP airspeed (significantly below VY) to allow compliance with the climb requirement of § 29.67(a)(1). Compliance with climb requirements can be substantiated initially by testing at a safe altitude above the ground. When OEI climb conditions are verified for Page B - 38

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weight, configuration, pressure altitude, and temperature, the CDP is then evaluated in a rejected takeoff.

(ii) A Category A takeoff procedure for which the CDP is defined as a specific “time,” “height,” or “airspeed” in the takeoff sequence combined with a precise takeoff crew procedure may be approved on the basis of equivalent safety when the following conditions can be satisfied:

(A) The flightcrew takeoff procedure must be shown to be consistently repeatable and not require exceptional piloting skill.

(B) It must be documented that the takeoff procedure will produce the required minimum energy level in terms of height and airspeed for all combinations of gross weight, altitude, and ambient temperature for which takeoff data are scheduled.

This may best be accomplished by conducting takeoff procedure abuse tests to show that variations from the established takeoff procedure that could reasonably be expected to occur in service do not result in significant increases in the takeoff distances.

(3) Rejected Takeoff Distance. The rejected takeoff is similar in many respects to the height-velocity (HV) tests described in paragraph AC 29.73. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at CDP are less critical than limiting HV points. As mentioned in paragraph AC 29.79, a minimum 5-knot clearance from any HV limiting condition should be provided throughout the takeoff flight path (see figure AC 29.63-1), and tests should be conducted simulating an unplanned engine cut. The HV diagram appropriate in the Category A test weights may be much less restrictive than that determined for Category B conditions. Normally, a minimum 1- second delay is applied after engine failure before pilot collective control corrections are allowed. However, if pilot cues are strong enough to make engine failure unmistakable, normal pilot reaction time may be utilized following engine failure. As in all engine failure testing, the pilot should not anticipate the failure by changing flight control positions or aircraft attitude. Average pilot techniques should be used. The two primary objectives of rejected takeoff testing are an assured capability to safely return to the takeoff surface when an engine fails at any point prior to CDP and the determination of the rejected takeoff distance that is needed when an engine fails at the CDP. It is important that the surface conditions be defined. For the rejected takeoff distance tests, a minimum of five satisfactory runs should be flown by the FAA/AUTHORITY pilot. The rejected takeoff distances from company and FAA/AUTHORITY runs may be averaged. The rejected takeoff distance tests will be used together with the OEI continued takeoff profiles to establish the required surface area for Category A operations.

(4) Continued Takeoff Distance.

(i) Continued takeoff profiles should be flown to determine the continued takeoff distance. This distance is measured from the point of takeoff initiation to the Page B - 39

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point in the takeoff profile where the following three conditions have all been attained after a failure of the critical engine at CDP: an airspeed equal to or greater than V TOSS, a positive rate of climb, and a height of at least 35 feet above the takeoff surface. The rotorcraft should not contact the ground at any point after engine failure. If the rotorcraft descends below 35 feet above the takeoff surface while accelerating to VTOSS, the takeoff distance is extended to the point that 35 feet is reattained with a positive rate of climb.

(ii) If the CDP is significantly above 35 feet so that the rotorcraft does not descend below 35 feet during acceleration to VTOSS, the takeoff distance then becomes the distance to the point in the takeoff profile at which both VTOSS and a positive rate-of-climb are attained after failure of the critical engine at CDP. For most applications, the rotorcraft should not be allowed to descend more than one-half the CDP height above the takeoff surface while accelerating to VTOSS. In addition, the rotorcraft should not be allowed to descend below the height above the takeoff surface at which a landing flare would normally be initiated. For example, if a rotorcraft has a CDP of 20 feet but when landing would normally initiate the landing flare at 15 feet, the takeoff profile should not be allowed to descend to 10 feet but should remain above 15 feet in establishing the takeoff distances.

(iii) In establishing the continued takeoff distance, the applicable pilot recognition delay time should be applied following the engine failure at CDP, and the takeoff profile should be established with the pilot using primary flight controls only to control the rotorcraft. The pilot engine failure recognition time delay before adjustment of the collective pitch control should be a minimum of 1 second unless it can be demonstrated that the pilot will have unmistakable engine failure cues sooner than 1 second.

(iv) Engine failure testing should be initially conducted at a safe distance above the ground to assess the continued takeoff profile before conducting the actual profiles for credit. This procedure will serve to validate predicted performance and may prevent an unexpected return to the surface during continued takeoff tests. A minimum of five acceptable runs should be flown by the FAA/AUTHORITY pilot, and these should be averaged with five acceptable runs flown by the manufacturer’s pilot.

(5) Abuse Testing. Takeoff procedure abuse tests should be conducted to show that reasonably expected variations in service from the established takeoff procedures do not result in a significant increase in the established takeoff distances.

Variations should include such considerations as under or over rotation during the takeoff initiation, under or over application of acceleration power, and missed CDP target parameters (e.g., time, height, or airspeed).

(6) Continued Climbout Path. The climb performance requirements of § 29.67(a)(1) should be met at the end of the continued takeoff distance segment.

Beginning at this point, the landing gear may be retracted, and secondary engine controls may be manipulated to adjust power. Any manipulation of secondary engine Page B - 40

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controls should be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls. The climb should be continued at VTOSS until approximately 100 feet above the takeoff surface. It should be demonstrated that the rotorcraft including any configuration changes can be accelerated from VTOSS to VY in a continuous maneuver without losing altitude. The airspeed and rotorcraft configuration (landing gear position, rotor RPM engine power, etc.) used to show compliance with the climb requirements of § 29.67(a)(2) should be attained at or prior to reaching 1,000 feet above the takeoff surface.

(7) Power. Power should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping to minimum specification values including consideration of temperature effects on engine power. Turbine engine power does not vary directly with density altitude (HD). At a given HD, turbine engine power available varies with ambient temperature. Turbine engines typically produce less horsepower as ambient temperature is increased (pressure altitude decreases) at a given density altitude, although some engines produce less horsepower at extremely cold temperatures. In either event, if one test sequence is to be utilized for a given HD, it would be appropriate to restrict test power to the lowest value attainable from a minimum specification engine through the approved ambient temperature range at the density altitude of the test. To attain maximum weights for varying ambient conditions, the applicant may utilize a parametric mapping of power available, pressure altitude, and temperature effects. For this case, engine topping may be adjusted throughout a range appropriate to the test HD.

(8) Aircraft Loading. Both forward and aft CG extremes should be spot checked to determine the critical loading for takeoff distances. Forward center of gravity is usually critical for continued takeoff distance tests while aft CG may be critical for the rejected takeoff because of over-the-nose visibility. A minimum of two weights should be flown at each altitude if the manufacturer elects to schedule field length variation as a function of gross weight. One weight should be the maximum weight for prevailing conditions and the other weight(s) should be low enough to attain a sufficient spread to verify weight accountability.

(9) Extrapolation. Weight cannot be extrapolated above test weight for the same reasons discussed in paragraph AC 29.79. See paragraph AC 29.45 regarding altitude extrapolation of test results.

(10) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data must be corrected for existing wind conditions during takeoff distance testing. Credit for headwind conditions may be given during flight manual data expansion. Refer to paragraph AC 29.45(b)(1) under “Winds for Testing” for allowable wind credit. Care should be applied in considering headwind credit for vertical operations as previous experience has resulted in difficulty collecting meaningful, repeatable data.

(11) Vertical Takeoffs.

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(i) General. Guidelines for rotorcraft certification using vertical takeoff techniques were developed and utilized for civil certification programs many years ago.

As experience has been gained, certain policy decisions have modified these guidelines. The following guidelines incorporate all available policy information as of January 1, 1981. The reader should be familiar with the preceding discussion regarding conventional Category A takeoff profiles because duplicate information is not repeated here.

(ii) Takeoff Profile. A typical vertical takeoff profile for a ground level heliport is shown in figure AC 29.59-1. The maneuver begins with the addition of sufficient power to initiate a climb to the CDP. It must be possible to make a safe landing without exceptional pilot skill if an engine fails at any point up to the CDP. At the CDP, the pilot becomes “go” oriented and continues the takeoff if an engine fails. A typical profile for pinnacle takeoff conditions is shown in figure AC 29.59-2.

Considerations are similar to those of the ground level heliport in figure AC 29.59-1;

however, the OEI pinnacle profile allows descent below the takeoff surface, specifies minimum edge clearance criteria, and allows relaxed requirements for final segment climb. Thus far, descent profiles up to 50 feet below the takeoff surface have been allowed; however, there is no reason why greater values could not be determined during engineering flight tests for certification. Use of such a profile, of course, would be dependent on obtaining an operational approval.

(iii) Critical Decision Point (CDP). For vertical takeoffs, the climb to CDP is nearly vertical, and CDP is typically defined primarily by height. Sufficient testing must be conducted to define a band of CDP conditions (heights) which will be consistent with anticipated variations in pilot technique and the minimum amount of equipment to be installed on the production aircraft. Rejected takeoffs are most critical from high CDPs, and continued OEI takeoffs are most critical from low heights. Tests at the extremes of this band are intended to verify that the anticipated CDP band is safe and repeatable in service for reasonable variations in pilot technique. These extreme points should not be used for distance determination when averaging takeoff performance data.

(iv) Conduct of the Test. Vertical takeoff profiles must be flown from a pad simulating operational conditions because the sight picture may be critical to successful OEI operations, particularly for elevated heliports. At all points on the vertical takeoff flight path up to the CDP, the pilot, with reasonable head movement, shall be able to keep sufficient portions of two heliport boundaries (front and one side) or equivalent markings in view to achieve a safe landing in case of engine failure.

Normally, a minimum 1-second delay is applied after engine failure before pilot collective control corrections are allowed. However, if pilot cues are strong enough to make engine failure unmistakable, normal pilot reaction time may be utilized following engine failure.

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(A) Establish the rejected takeoff distance as the horizontal distance from the rearmost point of the rotorcraft at the initiation of takeoff to the foremost point after the rotorcraft comes to a stop on the takeoff surface (including rotor tip path), assuming an engine failure in the vertical climb at the CDP; or (B) Establish the continued takeoff distance as the horizontal distance from lift-off to the point at which, following engine failure at CDP, the rotorcraft achieves 35 feet above the takeoff surface and VTOSS in a climbing posture. The continued takeoff profile from elevated heliports must clear the heliport obstructions by at least 15 feet vertically and 35 feet horizontally.

(v) Climb Requirements.

(A) The OEI takeoff profile should include a climb at VTOSS to 200 feet above the takeoff surface prior to accelerating to a higher speed.

(B) For elevated heliports, the climb requirement of § 29.67(a)(2) may be met at 200 feet above the takeoff surface or 1,000 feet above the surrounding terrain, whichever is higher.

(vi) Extrapolation. Basic guidelines for extrapolation are contained in paragraph AC 29.45. If, however, vertical takeoff weights are based upon allowable weights for hovering out-of-ground effect (OGE) with one engine inoperative, all vertical takeoff performance aspects may be extrapolated to the highest altitude requested for takeoff and landing.

(12) Night Operations.

(i) A minimum of three normal takeoffs (and landings) should be conducted to assure that aircraft lighting (internal and external) is adequate to allow normal Category A operations at night.

(ii) Engine failures should be simulated from points along the recommended takeoff profile. Night OEI rejected takeoffs and continued takeoffs from the CDP should be conducted to assure adequate night field of view and realization of Category A field lengths.

(iii) If special airfield markings are used as a reference or to define the CDP, the aircraft external lighting should be evaluated to assure that these airfield markings are adequately visible for night operations.

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Accelerate to Vv All Engine Climbout when clear of obstacles Accelerate to Vross CDP--+ Vertical Climb Rejected Takeoff Path Fl GURE AC 29.59-1 CATEGORY A VERTICAL TAKE OFF PROFILE

GROUND LEVEL HELIPORT AC 29-2C 9/30/99 Page B - 44

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Vertical Climb & t Rejected I Takeoff Path T t ~ Accelerate To Vross ✓ Minimum 15 Feet _______ ...._ ........ +--Minimum ____.

35 Feet All Engine Climbout OEI Flight Path FIGURE AC 29.59-2 CATEGORY A VERTICAL TAKEOFF PROFILE PINNACLE 9/30/99 AC 29-2C Page B - 45

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AC 29.59A (AC's 29.60, 29.61, & 29.62) §§ 29.59 (29.60, 29.61 and 29.62) (Amendment 29-39) TAKEOFF PATH, DISTANCE AND REJECTED TAKEOFF; GROUND LEVEL AND ELEVATED HELIPORT: CATEGORY A (For § 29.59 prior to Amendment 39, see paragraph AC 29.59.)

a. Explanation. Amendment 29- 39 moved the rejected takeoff requirements from § 29.55 to a new § 29.62 and clearly defined the takeoff path. It also added new §§ 29.60 and 29.61 to introduce the requirements for elevated heliport takeoff path, Category A and to more clearly define the parameters to be used in determining takeoff distance, respectively.

(1) Takeoff Decision Point. The Category A concept limits the rotorcraft takeoff weight such that if an engine failure is recognized at or before the TDP, a safe landing can be made or if an engine failure is recognized at or after the TDP, the takeoff can be continued. The purpose of these tests is to define the TDP, evaluate the necessary pilot techniques, and determine the required takeoff area for either alternative. The condition of equal distances for either stopping or continuing the takeoff is called a “balanced” field length. The combination of altitude and speed at the TDP which produces a balanced field length is not required for certification. This section deals with the Category A takeoff and rejected takeoff profiles. The profiles necessarily involve consideration of an average pilot skill level as well as a sequence in which it is assumed various configuration adjustments are made to the rotorcraft.

(2) Takeoff. The Category A takeoff path begins with an all -engines-operating acceleration segment to the engine failure point and continues with a one-engine-inoperative acceleration through the TDP to the takeoff safety speed (VTOSS). The engine failure point (EFP) and TDP are separated by pilot recognition time. (See Conventional Takeoff Profile, figure AC 29.53A-1, paragraph AC 29.53A of this advisory circular.) TDP is a “go/no- go condition which is analogous to V1 speed in transport airplanes. Prior to TDP the pilot is “stop” oriented, and when an engine failure is recognized in this portion of the takeoff, the pilot will abort because the rotorcraft has not yet achieved sufficient energy to assure continued flight. At the TDP the pilot becomes “go” oriented and when an engine failure is recognized at or beyond this point, the pilot will continue the takeoff because sufficient surface area no longer remains for an aborted takeoff. The takeoff flight path and the TDP should be defined such that a safe landing can be made from any point up to the TDP. This profile may differ significantly from the takeoff flight path developed for Category B weights. The TDP is the last point in the takeoff profile at which a rejected takeoff capability within the scheduled takeoff surface distance is assured. If an engine failure does not occur, the pilot continues the climb and accelerates past the TDP to the recommended climb speed.

(3) Rejected Takeoff. The rejected takeoff profile begins with an all engine acceleration segment to the EFP and ends when the rotorcraft is brought to a complete stop on the designated takeoff surface. The critical engine is made inoperative prior to Page B - 46

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the TDP, and the landing should be made with the remaining engine(s) operating within approved limits. The rejected takeoff distance is normally measured at a given reference point on the rotorcraft from the start of the takeoff to the same reference point after the rotorcraft has come to a complete stop. This distance should be increased by the rotorcraft length (including main and tail rotor tip paths).

(4) Takeoff Path.

(i) The transition to OEI flight takes place between the engine failure point and the point at which VTOSS is achieved. It is assumed that the maximum approved OEI power is used until the allowable time duration for that power is exhausted. It should be possible for the crew to fly the rotorcraft to V TOSS and attain an altitude of 35 feet and positive rate of climb and then climb to 200 feet above the takeoff surface or the lowest point in the takeoff path by flying the rotorcraft solely by the primary flight controls (including collective). At no time during the takeoff shall the rotorcraft descend below 15 feet above the takeoff surface when the TDP is above 15 feet. The landing gear may be retracted after attaining a speed of V TOSS, and a positive rate of climb. Flight manual procedures may recommend adjustment of auxiliary controls to improve OEI performance, but compliance with the performance requirements of § 29.67(a)(1) may not be based on use of secondary engine controls such as RPM beep switches. During the continued takeoff profile, the pilot is assumed to be flying the rotorcraft via the primary flight controls (cyclic stick, collective, and directional pedals). Manipulation of the throttle controls or beep switches may be permitted as long as such manipulation can be accomplished readily by the pilot flying the rotorcraft without removing his hands from the cyclic and collective flight controls.

These manipulations of engine controls should not make major adjustments in power and should not occur before attaining VTOSS. In no case should this be less than 3 seconds after the critical engine is made inoperative. The failure of one engine cannot affect continued safe operation of the remaining engines or require any immediate action by the crew per § 29.903(b). If a 30- second/2-minute or a 2 ½-minute power rating is used, it should be possible to complete the Category A takeoff profile (assuming recognition of an engine failure at or prior to the TDP), accelerate to VTOSS, attain 35 feet above the surface, stabilize in a climb of at least 100 feet per minute, and complete landing gear retraction prior to exhausting the 2 ½-minute time limit.

(ii) The takeoff safety speed, VTOSS, is a speed at which 100 FPM rate of climb is assured under conditions defined in § 29.67(a)(1). The takeoff distance is the distance from the start of the takeoff to the point at which VTOSS, 35 feet above the takeoff surface, and a positive rate of climb are attained.

(5) Continued Climbout Path. Continued acceleration and climb capability are assured by the 100 FPM VTOSS climb requirement of § 29.67(a)(1) and the 150 FPM requirement of § 29.67(a)(2), normally demonstrated at Vy. It sh ould be shown that the rotorcraft can be accelerated from VTOSS to VY in a continuous maneuver without losing altitude, including any configurative change (landing gear retraction, etc.). The distance Page B - 47

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required to accelerate from VTOSS to VY must be considered in determination of the climb and gradients required by § 29.1587(a)(6)(i) and (a)(6)(ii).

b. Procedures.

(1) Instrumentation. A photo theodolite, grid camera, GPS, or other position measuring equipment is normally required together with a ground station to measure wind, OAT, humidity (if applicable), and a two-way communication system to coordinate activities with the aircraft. A crash recovery team with support of a fire engine is highly desirable. Aircraft instrumentation should record with a time scale: engine parameters (speed, temperature, and power), rotor speed, flight parameters (airspeed, altitude, and normal acceleration as a minimum), flight control positions, power lever position, and landing gear loads. Additionally, a method should be devised to allow correlation of the aircraft instrumentation data with the space position data to accurately determine the length of the various takeoff segments.

(2) Establishing the Takeoff Decision Point (TDP).

(i) The TDP should be definable with the minimum crew using standard cockpit instrumentation. If a radar altimeter is used, it should be included in the minimum equipment list. If barometric altitude is used to define TDP, the operating conditions at which the altimeter is set should be defined. This is normally done on the ground with the minimum collective pitch. If the wind influences the altimeter reading, the correct relative wind information should be provided. Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the TDP becomes largely a function of the surface area required for takeoff. If takeoff conditions scheduled include considerable surface area (on the order of 2,000 feet), the TDP airspeed may be a high value near VY. This will allow a higher takeoff weight and demonstrate compliance with the VTOSS climb requirement of § 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If required surface area is a small value, TDP will necessarily be some lower airspeed value to allow for an aborted takeoff on the available surface. Weight may need to be reduced at lower values of TDP airspeed (significantly below VY) to allow compliance with the climb requirement of § 29.67(a)(1). Compliance with climb requirements can be substantiated initially by testing at a safe altitude above the ground. When OEI climb conditions are verified for weight, configuration, pressure altitude, and temperature, the TDP is then evaluated in a rejected takeoff.

(ii) A Category A takeoff procedure should satisfy the following conditions:

(A) The flightcrew takeoff procedure should be shown to be consistently repeatable and not require exceptional piloting skill.

(B) It should be documented that the takeoff procedure will produce the required minimum energy level in terms of height and airspeed for all combinations of Page B - 48

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gross weight, altitude, and ambient temperature for which takeoff data are scheduled.

This may best be accomplished by conducting takeoff procedure abuse tests to show that variations from the established takeoff procedure that could reasonably be expected to occur in service do not result in significant increases in the takeoff distances.

(3) Rejected Takeoff Distance. The rejected takeoff is similar in many respects to the height-velocity (HV) tests described in paragraph AC 29.73. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at TDP are less critical than limiting HV points. As mentioned in paragraph AC 29.79, a minimum 5-knot clearance from any HV limiting condition should be provided throughout the takeoff flight path (see figure AC 29.63-1), and tests should be conducted simulating an unplanned engine cut. The HV diagram appropriate to the Category A test weights may be much less restrictive than that determined for Category B conditions. Normally, a minimum 1- second delay (or pilot reaction time, whichever is greater) is applied after engine failure recognition, before pilot collective control corrections are allowed. If the rotorcraft incorporates an engine failure warning device, engine failure recognition should not be less than the time required for the engine to spool down and activate the device. As in all engine failure testing, the pilot should not anticipate the failure by changing flight control positions or aircraft attitude.

Average pilot techniques should be used. The two primary objectives of rejected takeoff testing are an assured capability to safely return to the takeoff surface when an engine failure is recognized at any point prior to TDP and the determination of the rejected takeoff distance required. It is important that the surface conditions be defined. The rejected takeoff distance tests will be used together with the OEI continued takeoff profiles to establish the required surface area for Category A operations.

(4) Takeoff Distance.

(i) Continued takeoff profiles should be flown to determine the continued takeoff distance. This distance is measured from the point of takeoff initiation to the point in the takeoff profile where the following three conditions have all been attained after a failure of the critical engine prior to TDP: an airspeed equal to or greater than VTOSS, a positive rate of climb, and a height of at least 35 feet above the takeoff surface.

If the rotorcraft descends below 35 feet above the takeoff surface while accelerating to VTOSS, the takeoff distance is extended to the point that 35 feet is reattained with a positive rate of climb.

(ii) If the TDP is significantly above 35 feet so that the rotorcraft does not descend below 35 feet during acceleration to VTOSS, the takeoff distance then becomes the distance to the point in the takeoff profile at which both VTOSS and a positive rate of climb are attained after failure of the critical engine prior to the TDP. For all applications, rotorcraft should not be allowed to descend below 15 feet above the takeoff surface while accelerating to VTOSS when TDP is above 15 feet. When TDP is below 15 feet, the aircraft should be able to accelerate in level flight or climb.

Fifteen feet should be considered the absolute minimum clearance allowed with greater Page B - 49

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AC 29-2C, Chg 2 4/25/06 clearances required for some rotorcraft dependent on rotorcraft geometry and performance characteristics. In addition, the rotorcraft should not be allowed to descend below the height above the takeoff surface at which a landing flare would normally be initiated. For example, a medium size twin- engined rotorcraft with a TDP of 100 feet or greater, using 20° nose down, would be expected to clear the ground by 25 feet whereas a large multiengined rotorcraft, using similar attitudes and TDP’s, would be expected to clear by 35 feet. For elevated heliports the rotorcraft may descend below the landing surface, but all parts of the rotorcraft must clear the heliport and all other obstacles by not less than 15 feet. These minimum heights would need to be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before and after TDP.

(iii) In establishing the continued takeoff distance, the applicable pilot recognition delay time should be applied following the engine failure prior to the TDP, and the takeoff profile should be established with the pilot using primary flight controls only to control the rotorcraft. The pilot engine failure recognition time delay before adjustment of the collective pitch control should be a minimum of 1 second.

(iv) Engine failure testing should be initially conducted at a safe distance above the ground to assess the continued takeoff profile before conducting the actual profiles for credit. This procedure will serve to validate predicted performance and may prevent an unexpected return to the surface during continued takeoff tests. A minimum of five acceptable runs should be flown by the FAA/AUTHORITY pilot, and these should be averaged with five acceptable runs flown by the manufacturer’s pilot.

(5) Abuse Testing. Takeoff procedure abuse tests should be conducted to show that reasonably expected variations in service from the established takeoff procedures do not result in a significant increase in the established takeoff distances.

Variations should include such considerations as under or over rotation during the takeoff initiation, under or over application of acceleration power, and missed TDP target parameters (e.g., time, height, or airspeed).

(6) Continued Climbout Path. The landing gear may be retracted at 35 feet.

The climb should be continued at VTOSS until 200 feet above the takeoff surface. The climb requirements of § 29.67(a)(1) should be met at 200 feet. It should be demonstrated that the rotorcraft, including any configuration changes, can be accelerated from VTOSS to VY in a continuous maneuver without losing altitude. The airspeed and rotorcraft configuration (landing gear position, rotor RPM engine power, etc.) used to show compliance with the climb requirements of § 29.67(a)(2) should be attained at or prior to reaching 1,000 feet above the takeoff surface.

(7) Power. Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping (maximum power available) to minimum specification values including consideration of temperature effects on engine power. If topping results in unrepresentative engine power management, the validity of the Cat A Page B - 50

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procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion a rejected take- off, it is acceptable for the engine and transmission transient range to be entered in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all conditions requested by the applicant. Any excursion beyond established transient limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.

(8) Turbine engine power does not vary directly with density altitude (HD). At a given HD, turbine engine power available varies with ambient temperature. Turbine engines typically produce less horsepower as ambient temperature is increased (pressure altitude decreases) at a given density altitude, although some engines produce less horsepower at extremely cold temperatures. In either event, if one test sequence is to be utilized for a given HD, it would be appropriate to restrict test power to the lowest value attainable from a minimum specification engine through the approved ambient temperature range at the density altitude of the test. To attain maximum weights for varying ambient conditions, the applicant may utilize a parametric mapping of power available, pressure altitude, and temperature effects. For this case, engine topping may be adjusted throughout a range appropriate to the test HD.

(9) Aircraft Loading. Both forward and aft CG extremes should be briefly checked to determine the critical loading for takeoff distances. Forward center of gravity is usually critical for continued takeoff distance tests while aft CG may be critical for the rejected takeoff due to forward and downward field of view. A minimum of two weights should be flown at each altitude if the manufacturer elects to schedule field length variation as a function of gross weight. One weight should be the maximum weight for prevailing conditions and the other weight(s) should be low enough to attain a sufficient spread to verify weight effect.

(10) Extrapolation. Takeoff and landing data may be extrapolated up to 4000 feet along an established W/σ line, to the maximum gross weight of the rotorcraft.

However, extrapolation will not be considered valid if unacceptable or marginally acceptable landing gear loads are experienced during testing at weights below the W/σ limit. See paragraph AC 29.77b(5) for further discussion of landing gear loads.

(11) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data should be corrected for existing wind conditions during takeoff distance testing. Credit for headwind conditions may be given during flight manual data expansion. Refer to paragraph AC 29.1587(a)(3)(iii) under “Wind Accountability” for allowable wind credit.

Care should be applied in considering headwind credit for vertical operations as previous experience has resulted in difficulty collecting meaningful, repeatable data.

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AC 29-2C, Chg 2 4/25/06

(12) Vertical Takeoffs.

(i) General. Guidelines for rotorcraft certification using vertical takeoff techniques were developed and utilized for civil certification programs many years ago.

As experience has been gained, certain policy decisions have modified these guidelines. The reader should be familiar with the preceding discussion regarding conventional Category A takeoff profiles because duplicate information is not repeated here.

(ii) Takeoff Profile. A ty pical vertical takeoff profile for a ground level heliport is shown in figure AC 29.59A-1. The maneuver begins with the addition of sufficient power to initiate a climb to the TDP. It should be possible to make a safe landing without exceptional pilot skill if an engine fails at any point up to the TDP less engine failure recognition time. At the TDP, the pilot becomes “go” oriented and continues the takeoff if an engine fails. The rotorcraft should not be allowed to descend below 15 feet above the takeoff surface during the continued takeoff. A typical profile for elevated heliports takeoff conditions is shown in figure AC 29.59A-2. Descent profile below the takeoff surface is allowed, after clearing the platform by at least a 15 feet radial margin, provided that the drop down height from the takeoff surface and the distance to reach VTOSS with a positive rate of climb is given in the performance chapter of the RFM.

(iii) Takeoff Decision Point (TDP). For vertical takeoffs, the climb to the TDP is nearly vertical, and the TDP is typically defined primarily by height. Sufficient testing should be conducted to define a band of TDP conditions (heights) which will be consistent with anticipated variations in pilot technique and the minimum amount of equipment to be installed on the production aircraft. Rejected takeoffs are most critical from high TDP’s, and continued OEI takeoffs are most critical from low heights. Tests at the extremes of this band are intended to verify that the anticipated TDP band is safe and repeatable in service for reasonable variations in pilot technique. These extreme points should not be used for distance determination when averaging takeoff performance data.

(iv) Conduct of the Test. Vertical takeoff profiles should be flown from a pad simulating operational conditions because the sight picture may be critical to successful OEI operations, particularly for elevated heliports. At all points on the vertical takeoff flight path up to the TDP, the pilot, with reasonable head movement, shall be able to keep sufficient portions of two heliport boundaries (front and one side) or equivalent markings in view to achieve a safe landing in case of engine failure.

Normally, a minimum 1-second delay or pilot recognition time interval, whichever is greater, is applied after the EFP before pilot collective control corrections are allowed. If the rotorcraft incorporates an engine failure warning device, engine failure recognition should not be less than the time required for the engine to spool down and activate the device.

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(A) Establish the rejected takeoff distance as the horizontal distance from the rearmost point of the rotorcraft at the initiation of takeoff to the foremost point after the rotorcraft comes to a stop on the takeoff surface (including rotor tip path), assuming an engine failure in the vertical climb at the TDP.

(B) Establish the continued takeoff distance as the horizontal distance from lift-off to the point at which, following engine failure prior to the TDP, the rotorcraft achieves; for a ground level heliport, 35 feet above the takeoff surface and VTOSS with a positive rate of climb; for an elevated heliport, the lowest point of the takeoff profile and not less than VTOSS with a positive rate of climb. The continued takeoff profile from elevated heliports should clear the heliport obstructions by at least a 15 feet radial margin.

(C) When used, the back-up technique usually requires the pilot to keep sufficient portions of the helipad in view and involves a rearward movement from the takeoff point to the TDP. In such cases the rearward horizontal distance required should be established as the distance from the rearmost point of the rotorcraft at the initiation of takeoff to the rearmost part of the rotorcraft at TDP. As stated in AC 29.45, crosswinds and tailwinds should be considered if requested by the applicant. Typically, this will require flight-testing to evaluate performance, pilot workload, field-of-view, and visual cueing.

(D) If special helipad markings or other non-standard external references are required to achieve the vertical takeoff performance, these special references should be included in the limitations section of the RFM.

(v) Climb Requirements.

(A) Ground level heliport. The OEI takeoff profile should include a climb at VTOSS to 200 feet above the takeoff surface then an acceleration in level flight from VTOSS to VY and a climb at VY to 1000 feet above the lowest point of the takeoff profile.

The climb requirements of § 29.67(a)(1) and (a)(2) may be met at referenced points located respectively at 200 feet and 1000 feet above the takeoff surface. The distance required to accelerate from VTOSS to VY must be considered in determination of the climb gradient required by § 29.1587 (a)(6)(i) and (a)(6)(ii).

(B) Elevated heliport. The OEI takeoff profile should include a climb at VTOSS to 200 feet above the lowest point of the takeoff profile then an acceleration in level flight from VTOSS to VY and a climb at VY to 1000 feet above the lowest point of the takeoff profile. The climb requirements of § 29.67(a)(1) and (a)(2) may be met at referenced points located respectively at 200 feet and 1000 feet above the lowest point of the takeoff profile.

(vi) Extrapolation. Basic guidelines for extrapolation are contained in paragraph AC 29.45. Weight can not be extrapolated above test weight. Altitude extrapolation should be limited to a maximum of ± 4000 feet.

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AC 29-2C, Chg 2 4/25/06

(13) Night Operations.

(i) A minimum of three normal takeoffs (and landings) should be conducted to ensure that aircraft lighting (internal and external) is adequate to allow normal Category A operations at night.

(ii) Engine failures should be simulated from points along the requested takeoff and landing profiles. Night OEI rejected takeoffs and continued takeoffs from the TDP and OEI landings from the LDP should be conducted at the requested WAT limiting conditions to ensure adequate night field of view, suitability of aircraft external lighting, and meets the Category A profiles.

(iii) If special airfield marking or lighting is used as a reference or to define the TDP, the aircraft external lighting should be evaluated to assure that the airfield marking or lighting is adequate for night operations.

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ACCELERATE TO Vy I/\IHEN CLEAR OF OBSTACLES TDP

**VERTICAL CLIMB ---- 15'**

**ALL ENGINE**

**CLIMBOUT**

ACCELERATE TO Vross ►

**OEI CLIMBOUT FIGURE AC 29.59A-1 CATEGORY A VERTICAL TAKEOFF PROFILE**

GROUND LEVEL HELIPORT 9/30/99 AC 29-2C Page B - 55

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TOP-~ VERTICAL i CLIMB &

**REJECTED**

**TAKEOFF PATH /**

**ALL ENGINE**

**CLIMBOUT**

ACCELERATE TO Vross

OEI FLIGHT PATH Takeoff distance FIGURE AC 29.59A-2 CATEGORY A VERTICAL TAKEOFF

PROFILE ELEVATED HELIPORT I Vross 1 and poslti11e 1 rate of climb I AC 29-2C 9/30/99 Page B - 56

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AC 29.63. § 29.63 (Amendment. 29-12) TAKEOFF: CATEGORY B.

a. Explanation.

(1) Takeoff distance is the horizontal distance measured from an initial position to a point 50 feet above the takeoff surface with all engines operating within approved limits.

(2) The height-velocity diagram is normally developed and accepted prior to conducting takeoff distance tests. Takeoff distance tests are conducted avoiding the critical areas of the diagram. The amount of power utilized in determining takeoff distance may not be greater than that used in constructing the takeoff corridor and “knee” portions of the height-velocity diagram. Power might also have to be constrained, depending upon the amount of excess power available, so that a “reasonable” nose down pitch attitude is not exceeded during the initial portion of the takeoff run. Acceptable values used during past programs include:

limits)

(i) Hover power + 10 percent (not to exceed rated engine takeoff power

(ii) A percent transmission limiting torque (not to exceed rated engine takeoff power limits), and

(iii) Engine (or transmission) limiting power for the particular ambient conditions.

(3) The critical center of gravity should be used for takeoff distance tests.

Critical center of gravity should be established analytically or from previous testing and may be forward or aft depending on the type of rotorcraft. Items that should be considered in determining the critical center of gravity are climb performance and cockpit visibility. At least two gross weights should be flown at each test altitude, if weight accountability is desired, in order to validate the manufacturers prediction of weight effects.

(4) The speed utilized at the 50-foot point in the takeoff profile (V50 speed) may be largely determined by the ability to obtain reliable, repeatable airspeed indications which can also comply with § 29.1323. Section 29.1323 ties the airspeed system accuracy requirements to the climbout speed. The climbout speed should be that speed attained at 50 feet in complying with § 29.63.

b. Procedures.

(1) Instrumentation. A ground station will measure ambient temperature, humidity (if applicable), and wind. For allowable wind conditions and engine power considerations refer to paragraph AC 29.45. A photo panel or hand recording method Page B - 57

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may be utilized, as necessary, to record engine and flight parameters. A phototheodolite, takeoff and landing camera, or other approved instrumentation is utilized to measure distance, heights, speed, and time.

(2) Conduct of the Test. If the applicant elects to show weight effects on distance, at least two weights should be flown and, depending on the range of takeoff and landing altitudes to be approved, at least two test altitudes should be flown.

Altitudes should be sufficiently far apart to include a major portion of the approved takeoff and landing altitude range. Takeoff profiles should be started from an initial condition. For takeoffs from a hover, the hover height should be determined by performing fixed collective takeoffs as described in paragraph AC 29.51. “Takeoff” power should be smoothly applied and the aircraft nose lowered as necessary to accelerate without gaining excessive altitude. It must be possible to conduct a consistent takeoff profile clear of the height-velocity diagram with normal pilot effort and skill. A minimum of five good runs should be flown by the FAA/AUTHORITY pilot at each altitude and weight. Runs by the company and FAA/AUTHORITY pilot may be averaged. Effects of missing the V 50 speed by some amount (± 5 knots, for example) or other small changes in profile should be evaluated to determine if gross performance changes result from small piloting errors. Engine failures should be conducted along the takeoff profile to assure safe landing capability. Past programs have shown the low speed point immediately after addition of power to be particularly critical. Night takeoffs should at least be qualitatively evaluated to assure the takeoff procedures are compatible for night operation.

(3) Test Results. Test results are utilized in constructing the flight manual takeoff distance charts required by § 29.1587. The takeoff surface utilized in conducting these takeoff distance and engine failure tests should be included in the flight manual.

The “climbout speed” should also be defined and included in the flight manual. The airspeed utilized at the 50-foot point in the conduct of these tests must be clearly defined to allow compliance with § 29.1323. Test results may be extrapolated in accordance with guidance contained in paragraph AC 29.45.

(4) Test Techniques. For the FAA/AUTHORITY test data runs which will result in rotorcraft flight manual (RFM) performance, only the operational cockpit instrumentation as shown on the minimum equipment list and the piloting procedures from the RFM should be used. A useful technique is to “lead” the targeted V 50 speed by a fixed amount, so that a smooth, consistent, and operationally realistic transition may be made between the acceleration and climbout phases; e.g., begin rotation at 35 knots to achieve 46 knots passing 50 feet. This and other pertinent information defining the takeoff flight path are required flight manual entries per § 29.1587(b).

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0 z :::)

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VELOCITY (KNOTS) Fl GURE AC 29.63-1 CONVENTIONAL TAKE OFF PROFILE CATEGORY B 9/30/99 AC 29-2C Page B - 59

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AC 29.65. § 29.65 (Amendment 29-15) CLIMB: (ALL ENGINES OPERATING).

a. Explanation.

(1) Section 29.65 requires in part that the steady rate of climb be determined for each Category B rotorcraft with maximum continuous power on each engine for the range of weights, altitudes, and temperatures for which certification is requested. The climb airspeed should be the best rate-of-climb (VY) for standard day sea level conditions at maximum weight and at a speed(s) selected by the applicant for other conditions not to exceed VNE. The applicant can either publish a climb schedule in accordance with the above or utilize a constant climb airspeed for all conditions.

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.

(2) For Category A rotorcraft, 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 the best rate-of-climb airspeed VY. If such a technique is used, climbs should be flown in pairs on opposite headings 90° to the winds at the test altitude. This procedure will minimize any windshear effects.

All testing should be done in smooth air. Windshear is usually an indication of unstable air or a temperature inversion and should be avoided. The climbs are flown on reciprocal headings for approximately 5 minutes through a 1,000-foot band, or a comparable time/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 sufficient to bracket 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 the best rate-of-climb airspeed (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 Page B - 60

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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 at approximately 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 installed minimum specification engine.

c. Procedure to Determine all Engine Operating Climb Performance.

(1) Background. Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to obtain the rotorcraft’s climb performance for the flight manual. 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 techniques used to determine this performance may be the same as those used in the VY determination. The climbs are conducted on reciprocal headings at the established airspeed(s) through the target altitude range.

The same parameters are recorded. 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 accurate. 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 29.65A (AC 29.64) §§ 29.64 and 29.65 (Amendment 29-39) CLIMB (GENERAL AND ALL ENGINES OPERATING).

a. Explanation.

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(1) Amendment 29-39 relocated and clarified the general climb requirements into a new § 29.64 and added requirements to determine Category A climb performance in § 29.65. The guidance material presented in paragraph AC 29.67 does not apply to rotorcraft certified with Amendment 29-39 or later. Sections 29.64 and 29.65 require that the steady rate of climb be determined with maximum continuous power on each engine for the range of weights, altitudes, and temperatures for which certification is requested. The climb airspeed should be the best rate-of-climb (VY) for standard day sea level conditions at maximum weight and at a speed(s) selected by the applicant for other conditions not to exceed VNE. The applicant can either publish a climb schedule in accordance with the above or utilize a constant climb airspeed for all conditions.

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 should 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.

(2) If VNE at any altitude is less than the maximum gross weight sea level standard day condition VY, the steady rate-of-climb should be determined at the climb speed(s) selected by the applicant not to exceed VNE. The climb performance should 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 the best rate-of-climb airspeed VY. If such a technique is used, climbs should be flown in pairs on opposite headings 90° to the winds at the test altitude. This procedure will minimize any windshear effects.

All testing should be done in smooth air. Windshear is usually an indication of unstable air or a temperature inversion and should be avoided. The climbs are flown on reciprocal headings for approximately 5 minutes through a 1,000-foot band, or a comparable time/altitude band, using maximum continuous power at a constant airspeed. Periodic power adjustments may be necessary. Additional reciprocal heading climbs should also be conducted at different airspeeds sufficient to bracket 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 the best rate-of-climb airspeed (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 Page B - 62

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started at the desired W/σ with maximum continuos 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 at approximately 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 result in the airspeed for maximum rate of climb. However, aircraft stability may suggest that a higher climb speed may be used for VY.

(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 installed minimum specification engine.

c. Procedure to Determine all Engine Operating Climb Performance.

(1) Background. Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to obtain the rotorcraft’s climb performance for the flight manual. 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 techniques used to determine this performance may be the same as those used in the VY determination. The climbs are conducted on reciprocal headings at the established airspeed(s) through the target altitude range.

The same parameters are recorded. The rotorcraft will usually climb very rapidly during the first few thousand feet; therefore, the data acquisition method should be timely if accurate results are expected. This procedure is usually repeated at weight extremes.

The resulting data should 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 accurate. 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 should 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 29.67. § 29.67 (Amendment 29-34) CLIMB: ONE ENGINE INOPERATIVE.

a. Explanation.

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(1) Section 29.67 requires that Category A rotorcraft must be capable of a steady rate-of-climb without ground effect, of at least 100 feet per minute for all combinations of weight, altitude, temperature, and center of gravity for which takeoffs are to be scheduled. The rate- of-climb is determined with the critical engine inoperative and the remaining engine(s) operating within approved operating limits. The landing gear is extended and the airspeed is the takeoff safety speed (VTOSS) selected by the applicant.

(2) In addition, the steady rate-of-climb must be at least 150 feet per minute at 1,000 feet above the takeoff surface for which takeoffs are to be scheduled. The rate-of-climb will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or the 30-minute minimum specification installed power available values. The landing gear is retracted and the airspeed is that selected by the applicant.

b. Procedures.

(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (paragraph AC 29.65) except that the VTOSS and the Category A climb speed may be selected by the applicant for different weights and ambient conditions. The Category A climb speed could be a single speed, vary as VY does, or actually be VY. Making a Category A climbout speed equal to VY should be encouraged to simplify cockpit procedures. The required results are the allowable weight, altitude, and temperature combinations wherein the rotorcraft is capable of demonstrating 100 feet per minute rate-of-climb at VTOSS and 150 feet per minute rate-of-climb at 1,000 feet above the takeoff surface. Either of these two climb requirements may establish the maximum allowable takeoff weight.

(2) For multiengine Category B rotorcraft with engine isolation, the steady rate of climb or descent must be determined at VY, using maximum continuous power and 30-minute power if that rating is approved. Appropriate performance data must be included in the Rotorcraft Flight Manual to cover variations in gross weight, altitude, and temperature.

(3) Since climb performance testing is normally conducted separately from Category A and B takeoff performance testing, it is imperative the engine power(s), rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb performance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilot/crew actions during the Category A takeoff and climbout maneuver must be thoroughly evaluated. The pilot’s full attention is required to control the rotorcraft during this phase of flight. Permitting the pilot to readjust (beep) the rotor RPM during this phase of flight should be considered only if such adjustment can be accomplished without a significant increase in pilot workload.

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(4) A typical sequence for selecting the various speeds to comply with this requirement is as follows:

(i) Conduct sawtooth climbs at the various airspeeds (VY and below) up to the proposed takeoff and landing altitudes. From this a determination can be made regarding the maximum allowable weight that will result in a rate of climb of 150 feet per minute at the selected VY for the proposed ambient conditions.

(ii) At the same time determine the minimum value of VTOSS that will result in 100 feet per minute rate of climb at the maximum weight determined in (b)(4)(i).

AC 29.67A. § 29.67 (Amendment 29-39) CLIMB: ONE ENGINE INOPERATIVE.

a. Explanation.

(1) Amendment 29-39 expanded the OEI rate of climb requirements. The guidance material presented in paragraph AC 29.67 does not apply to rotorcraft certified with Amendment 29-39 or later. Section 29.67 requires that Category A rotorcraft should be capable of a steady rate-of-climb without ground effect 200 feet above the takeoff surface, of at least 100 feet per minute for all combinations of weight, altitude, temperature, and center of gravity for which takeoffs are to be scheduled. The rate-of-climb is determined with the critical engine inoperative and the remaining engine(s) operating within approved operating limits. The landing gear is extended and the airspeed is the takeoff safety speed (VTOSS) selected by the applicant.

(2) The steady rate-of-climb should be at least 150 feet per minute at 1,000 feet above the takeoff surface for which takeoffs are to be scheduled. The rate- of-climb will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or the 30-minute minimum specification installed power available values. The landing gear is retracted and the airspeed is that selected by the applicant.

(3) Additionally, the steady state rate of climb or descent must be determined with the critical engine inoperative and the remaining engines at OEI maximum continuous power and at 30-minute OEI power if applicable. This performance must be scheduled throughout the ranges of weight, altitude and temperatures for which certification is requested with the landing gear retracted, at an airspeed selected by the applicant.

b. Procedures.

(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (paragraph AC 29.65) except that the VTOSS and the Category A climb speed may be selected by the applicant for different weights and ambient conditions. The Category A climb speed could be a single speed, vary as VY does, or actually be VY. Making a Category A climbout speed equal to VY should be encouraged to simplify cockpit Page B - 65

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procedures. The required results are the allowable weight, altitude, and temperature combinations wherein the rotorcraft is capable of demonstrating 100 feet per minute rate-of-climb at VTOSS at a height of 200 feet above the takeoff surface and 150 feet per minute rate-of-climb at 1,000 feet above the takeoff surface. Either of these two climb requirements may establish the maximum allowable takeoff weight.

(2) For multiengine Category B rotorcraft with engine isolation, the steady rate of climb or descent should be determined at VY, using maximum continuous power, maximum continuous OEI power, and 30-minute power if that rating is approved.

Appropriate performance data should be included in the Rotorcraft Flight Manual to cover variations in gross weight, altitude, and temperature.

(3) Since climb performance testing is normally conducted separately from Category A and B takeoff performance testing, it is imperative the engine power(s), rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb performance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilot/crew actions during the Category A takeoff and climbout maneuver should be thoroughly evaluated. The pilot’s full attention is required to control the rotorcraft during this phase of flight. Permitting the pilot to readjust (beep) the rotor RPM during this phase of flight should be considered only if such adjustment can be accomplished without a significant increase in pilot workload.

(4) A typical sequence for selecting the various speeds to comply with this requirement is as follows:

(i) Conduct sawtooth climbs at the various airspeeds (VY and below) up to the proposed takeoff and landing altitudes. From this, a determination can be made regarding the maximum allowable weight that will result in a rate of climb of 150 feet per minute at the selected VY for the proposed ambient conditions.

(ii) At the same time, determine the minimum value of VTOSS that will result in 100 feet per minute rate of climb at the maximum weight determined in b(4)i.

AC 29.71. § 29.71 (Amendment 29-12) ROTORCRAFT ANGLE OF GLIDE:

**CATEGORY B.**

a. Explanation.

(1) Performance capabilities during stabilized autorotative descent are useful pilot tools to assist in the management of a Category B rotorcraft 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 are required as flight manual entries per § 29.1587. The speed for minimum rate of descent is useful for Page B - 66

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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. This speed, with appropriate rotor RPM, provides 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 29.71-1.

(4) Forward center of gravity is usually critical, however, center of gravity effects should be spot-checked to confirm this for a given design.

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) 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.

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**RATE OF**

**DESCENT RECOMMENDED**

AUTOROTATIONAL SPEED ~ I

**MINIMUM /**

**RATE OF DESCENT / / / / / / / /**

AIRSPEED ~

BEST GLIDE ANGLE Fl GURE AC 29. 71-1 AUTO ROTATIONAL CHARACTER! STI CS - TYPICAL AC 29-2C 9/30/99 Page B - 68

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AC 29.73. § 29.73 (Amendment 29-3) PERFORMANCE AT MINIMUM OPERATING SPEED. HOVER PERFORMANCE FOR ROTORCRAFT .

(For performance at minimum operating speed and for hover performance after Amendment 38, see § 29.49 and paragraph AC 29.49).

a. Explanation.

(1) For the purpose of this manual, 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.

(2) The regulatory requirement for hover performance, § 29.73, 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 weights. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle 1 ½ 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, a minimum amount of data are 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 landings. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.

(5) The minimum hover height 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 29.51 for the procedure to determine the minimum allowable hover height.

b. Procedures.

(1) Two methods of acquiring hover performance data are the tethered and 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 Page B - 69

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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 stabilized point. The total thrust the rotor produces is the rotorcraft’s gross weight, 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 s ome 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 cable and load cell are perpendicular to the ground. To insure 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. This can also be accomplished by attaching two accelerometers to the load cell which sense 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 or less as there are no accurate methods of correcting hover data for wind effects. Large rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are:

(i) Engine torques.

(ii) Rotor speed.

(iii) Ambient temperatures.

(iv) Pressure altitude.

(v) Fuel used (or remaining).

(vi) Load cell reading.

(vii) Generator(s) load.

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 Page B - 70

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data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in paragraph b(4) below 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 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 is obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet.

AC 29.75. § 29.75 (Amendment 29-17) LANDING.

a. Explanation.

(1) This rule incorporates all of the landing performance requirements for transport category rotorcraft. It consolidates requirements for landing data, Category A landing, Category A flight data, and Category B landing. Parallel takeoff requirements are located in four separate sections of the rule, §§ 29.51 through 29.63. As such, to assure necessary subjects are treated separately, the following discussion will be separated into three parts: (a) a general discussion of basic landing distance requirements, (b) Category A requirements (including vertical landing), and

(c) Category B requirements.

(2) All landing performance data are corrected to a smooth, dry, hard, level landing surface condition. As with other flight maneuvers, landings must be accomplished with acceptable flight and ground characteristics using normal pilot skills.

The rule states that Category A and B landing data must be determined at each approved WAT (Weight, Altitude, Temperature) condition. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 29.45. As in other performance areas, engines must be Page B - 71

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operated within approved limits. Power considerations are the same as those described under paragraph b(2)(ii)(C).

(3) Unlike fixed wing aircraft, rotorcraft typically require significantly more landing surface area with an engine inoperative than with all engines operating.

Because of this characteristic, the landing distance requirements are met with at least one engine inoperative to assure the most conservative landing distance measurement is achieved.

b. Procedures.

(1) Category A Requirements.

(i) Explanation. The Category A certification concept limits landing weight to a value that will allow the rotorcraft, following an engine failure at the landing decision point (LDP), to land within the available runway or to execute a balked landing, descending no lower than 35 feet above the landing surface. See figure AC 29.75-1.

(A) LDP. The Category A landing profile begins with an assumed engine failure at or prior to the LDP. The LDP is typically defined in terms of airspeed, rate of descent, and altitude above the landing surface. The approach path angle can be defined by LDP airspeed and rate of descent values. Definition of the LDP should include an approach angle because both the landing distance and the missed approach path are significantly influenced by landing approach angle. At any point in the single engine approach path down to and including the LDP, the pilot may elect to land or to execute a balked landing and he is assured both an adequate surface area for OEI landing and adequate climb capability for an OEI balked landing. Said another way, if an engine fails at any point down to and including the LDP, the pilot may safely elect to land or to “go around” by executing a balked landing. The LDP must be defined to permit acceleration to VTOSS at an altitude no lower than 35 feet above the landing surface. The LDP represents a “commit” point for landing. Prior to the LDP in the one engine inoperative approach, the pilot has a choice, he may either land or fly away.

After passing the LDP he no longer has sufficient energy to assure transition to a balked landing condition without contacting the landing surface. If an engine fails after LDP in a normal (all engine) landing the pilot is committed to land. The LDP and landing approach path must be defined such that the critical areas of the height-velocity diagram are avoided. A typical LDP for conventional Category A profiles is 100 feet above the landing surface. LDP should be specified in terms of both actual altitude above the landing surface and indicated barometric altitude. Speed at the LDP should be specified in terms of indicated airspeed.

(B) Landing distance. Approach and landing path requirements are stated in general terms in paragraphs (b)(2) and (4) of § 29.75. The approach path must allow smooth transition for one engine inoperative landing and for balked landing maneuvers and must allow adequate clearance from potentially hazardous HV combinations. Paragraph (b)(4)(ii) implies that a less restrictive HV envelope may exist Page B - 72

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for the Category A approach condition in comparison to that determined under high power conditions in § 29.79. The manufacturer may elect to use this added capability.

The added capability arises from the fact that lower power levels, a lower collective setting, and an established rate of descent accompany typical approach conditions as opposed to the more critical high power conditions of § 29.79. Landing distance is measured from a point 50 feet (25 feet for VTOL) above the landing surface to a stop.

For flight manual purposes, the distance is from the point at which the lowest part of the rotorcraft first reaches 50 feet (25 for VTOL) to the foremost point of the rotorcraft (including rotor tip path) after coming to a stop.

(C) All engine out landing. Section 29.75(b)(5) contains the Category A certification requirement for “last” engine failure and all engine inoperative landing. The rule states that it must be possible to make a safe landing on a prepared surface after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously. See paragraph AC 29.143a(2)(iii)(A) for the Category A sequential engine failure criteria. The conditions for last engine failure are maximum continuous power or 30-minute power if that rating is approved, “wings” level flight, and sudden engine failure with a pilot delay of 1 second or normal pilot recognition time, whichever is greater. Complete power failure has occurred in twin engine Category A rotorcraft.

This requirement ensures 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 V NE engine failure testing where instrumentation and critical loading have been established for those test conditions. See discussion under paragraph AC 29.143. The all engine out landing is ordinarily conducted in conjunction with an HV or Category A landing distance phase where ground instrumentation and safety equipment are available. The rotorcraft must be capable of conducting the all engine out landing at the takeoff and landing WAT limiting conditions up to the maximum altitude approved for takeoff and landing.

(ii) Procedures.

(A) Instrumentation/Equipment. Instrumentation requirements are basically the same as those for Category A takeoff. A photo theodolite, grid camera, or other position measuring equipment is needed, along with a ground station to measure wind, OAT, and humidity (if applicable). A two- way communication system between the aircraft and the position measuring equipment is essential. Aircraft instrumentation should include engine and flight parameters, control positions, power lever position, landing gear loads, and a method for synchronizing power cuts between the external light normally used for photo theodolite or camera, and onboard instrumentation. 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 support of a fire engine is highly desirable.

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(B) Establishing the LDP.

(1) Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the LDP becomes largely a function of the runway length required for landing. If landing conditions to be scheduled include considerable runway length (on the order of 1,000 feet) the LDP may be defined at a relatively high speed allowing transition to a takeoff safety speed near VY which will allow the maximum amount of weight for compliance with the balked landing climb requirements of § 29.77(b)/§ 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If the runway length is small, LDP will typically be at a lower speed and may be at a higher altitude to allow balked landing transition within the available distance. Landing weight may need to be reduced to allow landing from the lower speed or higher altitude decision point for shorter landing distances. In this case the requirements of § 29.67(a)(1) may be limiting. The climb performance and climb speeds required by § 29.67(a)(1) and (2) should be established prior to Category A landing tests.

(2) The one engine inoperative landing is similar in many respects to the height-velocity tests described in paragraph AC 29.79. Most of the comments, cautions, and techniques for HV also apply here even though typical flight conditions at LDP 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 and one engine is made inoperative prior to LDP After the LDP, speed is reduced and the rotorcraft is flared to a conventional one engine inoperative landing. Depending on the landing characteristics and landing profile, the flare may be initiated either prior or subsequent to the 50-foot elevation utilized in determining landing distance. Testing should include an engine failure at the LDP with a 1-second pilot delay to assure safe landing capability for this critical case. A minimum of five acceptable runs for distance should be flown by the FAA/AUTHORITY pilot. These may be averaged with an equal number of acceptable runs by the company pilot.

(3) The balked landing portion of the landing profile is addressed under § 29.77, Balked Landing: Category A. For an explanation of that requirement and a discussion of those test procedures refer to paragraph AC 29.71.

(C) 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 sy stem it becomes vitally important to check topping power prior to each test sequence.

(D) Aircraft Loading. Aft center of gravity is usually most critical for landing distance determination because visibility constraints limit the degree to which the pilot can flare the rotorcraft for landing. If a weight effect is shown, a minimum of Page B - 74

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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.

(E) Extrapolation. Weight cannot be extrapolated above test weight. See discussion under Height-Velocity Testing in paragraph AC 29.79. If no marginal areas are apparent and an acceptable analytical method is used, performance data may be extrapolated ±4,000 feet density altitude from test conditions. (See paragraph AC 29.45.)

(F) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data must be corrected for existing wind conditions during landing distance tests. Credit for headwind conditions may be given during flight manual data expansion. Paragraph AC 29.45 details allowable wind credit.

(G) All engine out landing.

(1) Several procedures can be utilized to demonstrate compliance with the all engine out landing requirement. As discussed in the explanation portion of this paragraph, § 29.75(b) contains two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. This requirement is discussed in paragraph AC 29.143. The second aspect of this rule requires that a landing from autorotation be possible on a prepared surface. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above a prepared landing surface. If a complete company test program has documented an all engine out landing to the GW/σ (gross weight/density ratio) limit for takeoff and landing at each altitude, 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.

(2) 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 assure a reasonable chance of survivability for the all engine failure condition. The touchdown speed (less than 40 KIAS is recommended) should be consistent with the type design limits including landing gear capability, aircraft visibility, and any other factors affecting repeatability of the maneuver. On Category A rotorcraft, rotor inertia is typically much lower than for single engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Also, 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 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 Page B - 75

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testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines cut off.

(3) 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 assuring 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 Category A rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (reference paragraph AC 29.79.) 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.

(H) Vertical Landings. The reader should be familiar with the preceding discussion of conventional Category A landing profiles because duplicate information is not repeated here. A typical vertical landing profile is shown in figure AC 29.75-2. This profile is equally applicable to both ground level and pinnacle sites. The profile begins at a stabilized single engine approach condition. It must be possible to make a safe OEI landing or go-around at any point prior to the LDP. At the LDP the aircraft becomes committed to landing. A safe landing must be possible in case of an engine failure at any point before or after the LDP. Testing should include a simulated failure at LDP with a 1-second delay or normal pilot response time, whichever is longer, and subsequent landing within the allowable area. The LDP is typically well above the 25-foot point from which landing distance is measured. The landing distance is the distance from the point at which the lowest portion of the rotorcraft reaches 25 feet above the landing surface to the forward-most point after coming to a stop (including main rotor tip path). The LDP becomes very important for landing on small, elevated heliports. The LDP must be clearly defined and flight manual instructions should carefully explain any pilot procedures. An illustration similar to figure AC 29.75-2 with somewhat more detailed information is most useful. Night OEI landings should be conducted to verify suitable visibility for both internal and external vertical landing cues.

c. Category B Requirements.

(1) Explanation. Section 29.75(c) contains the Category B landing requirements. For rotorcraft that do not meet the Category A powerplant installation requirements of this part, landing tests are conducted with all engines inoperative in an autorotative descent condition. Landing distance is measured from the 50- foot point to the point at which the rotorcraft is completely stopped (approximately 3 knots for water Page B - 76

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landings). The autorotative approach speed is selected by the applicant. The landing maneuver is similar to that referred to during normal training flights as a practice autorotation. As in HV tests, care must be taken to assure no power is delivered to the rotor drive system during these tests. A small amount of power can have a significant effect on landing test results. Multiengine rotorcraft incorporating Category A engine isolation features may conduct landing distance tests with only one engine inoperative using the procedures prescribed above for Category A. For these rotorcraft the one engine inoperative condition typically results in much shorter distances due both to a much lower speed at the 50-foot point and the added power available for flaring and cushioning the landing. Instrumentation requirements are the same as those described under Category A above. Appropriate ambient conditions and allowable extrapolation are discussed under paragraph AC 29.45.

(2) Procedures. Prior to conducting these tests the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. For Category B rotorcraft without engine isolation, 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 landing flare may be initiated prior to the 50- foot point. 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 flare-ability. Following the flare, the rotorcraft is allowed to touchdown in a landing attitude. Rotor RPM at touchdown should be recorded and it must be within allowable structural limits. For wheeled rotorcraft, the brakes are applied to an incipient skid for most efficient stopping. For rotorcraft on skids, the collective should be lowered as soon as characteristics allow in order to place a greater weight on the landing skids.

These procedures would be appropriate flight manual entries to show how landing distances can be realized. For flight manual purposes the landing distance should include the horizontal distance from the point at which the lowest part of the rotorcraft first reaches 50 feet above the landing surface to the point at the foremost part of the rotorcraft (including rotor tip path) after coming to a stop. For Category B rotorcraft with engine isolation, the landing procedures are as described for Category A landing.

When conducting Category B landings utilizing Category A “procedures,” § 29.75(b)(2) can be misleading. No transition capability to balked landing is intended for Category B rotorcraft. Section 29.77, Balked Landing, Category A, applies only to Category A rotorcraft and not to Category B rotorcraft which incorporates Category A “design” features. Five acceptable landing runs should be flown by the FAA/AUTHORITY pilot at each test weight. Results may be averaged with an equal number of company runs. If a weight effect on landing distance is to be shown, a minimum of two weight extremes are normally tested.

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Vross Vy Vross 1-<1---- LANDING DISTANCE--• FIGURE AC 29.75-1 CATEGORY A CONVENTIONAL LANDING - CLEAR HELIPORT AC 29-2C 9/30/99 Page B - 78

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FIGURE AC 29.75-2 CATEGORY A VERTICAL LANDING Page B - 79

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AC 29-2C, Chg 4 5/1/2014 AC 29.75A. (AC's 29.77, 29.79, 29.81, & 29.83) §§ 29.75, 29.77,

29.79, 29.81, and 29.83 (Amendment 29-39) LANDING.

(For § 29.77 and § 29.79 prior to Amendment 29-39, see paragraphs AC 29.77 and AC 29.79 respectively.)

a. Explanation.

(1) Amendment 29-39 revised and relocated many of the landing requirements of Part 29. Changes were made to the general landing requirements of § 29.75. New requirements were added for designating a landing decision point (LDP) in § 29.77.

The original § 29.79 was redesignated as a new § 29.87. Category A landing requirements were established in a new § 29.79. Requirements were added to determine landing distances in a new § 29.81. Revised Category B landing requirements were relocated from § 29.75(c) into a new § 29.83. The guidance material from paragraph AC 29.75 does not apply to rotorcraft certified with Amendment 29-39 or later.

(2) These rules incorporate all of the landing performance requirements for transport category rotorcraft. They contain the requirements for landing data, Category A landing, and Category B landing. Parallel takeoff requirements are located in eight separate sections of the rule, §§ 29.51 through 29.63. As such, to ensure that necessary subjects are treated separately, the following discussion will be separated into three parts: (a) a general discussion of basic landing distance requirements,

(b) Category A requirements (including vertical landing), and (c) Category B requirements.

(3) All landing performance data are corrected to a smooth, dry, hard, level landing surface condition. As with other flight maneuvers, landings should be accomplished with acceptable flight and ground characteristics using normal pilot skills.

The rule states that Category A and B landing data should be determined at each approved WAT (Weight, Altitude, Temperature) condition. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 29.45. As in other performance areas, engines should be operated within approved limits. Power considerations are the same as those described under paragraph b(1)(ii)(C).

(4) Unlike fixed-wing aircraft, rotorcraft typically require significantly more landing surface area with an engine inoperative than with all engines operating.

Because of this characteristic, the Category A landing distance requirements are met with at least one engine inoperative to ensure the most conservative landing distance measurement is achieved.

b. Procedures - Category A Requirements.

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(1) Explanation. The Category A certification concept limits landing weight to a value that will allow the rotorcraft, following an engine failure at the landing decision point (LDP), to land within the available area or to execute a balked landing descending no lower than 15 feet (or higher depending on rotorcraft geometry and performance characteristics) above the landing surface. For elevated heliports the rotorcraft may descend below the landing surface, but all parts of the rotorcraft must clear the heliport and other obstacles by not less than 15 feet. These minimum heights should be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before and after the LDP. See figure AC 29.75A-1. For additional information addressing the OEI landing case at night, refer to AC 29.59A.b.(13).

(i) LDP. The Category A landing profile begins with an assumed engine failure at or prior to the LDP. The LDP is typically defined in terms of airspeed, rate of descent, and altitude above the landing surface. The approach path angle can be defined by LDP airspeed and rate of descent values. Definition of the LDP should include an approach angle because both the landing distance and the missed approach path are significantly influenced by landing approach angle. At any point in the single engine approach path down to and including the LDP, the pilot may elect to land or to execute a balked landing and he is assured both an adequate surface area for OEI landing and adequate climb capability for an OEI balked landing. Said another way, if an engine failure is recognized at any point down to and including the LDP, the pilot may safely elect to land or to “go-around” by executing a balked landing. The LDP should be defined to permit acceleration to VTOSS clearing the landing surface by a minimum of 15 feet. The LDP represents a “commit” point for landing. Prior to the LDP in the one engine inoperative approach, the pilot has a choice, he may either land or fly away. After passing the LDP, he no longer has sufficient energy to assure transition to a balked landing condition without contacting the landing surface. If an engine failure is recognized after LDP in a normal (all engine) landing, the pilot is committed to land.

The LDP and landing approach path should be defined such that critical areas of the height-velocity diagram are avoided. A typical LDP for conventional Category A profiles is 100 feet above the landing surface. LDP should be specified in terms of both actual height above the landing surface and indicated barometric altitude. Speed at the LDP should be specified in terms of indicated airspeed. The applicant may elect to develop an alternate all-engines-operating (AEO) approach procedure which meets the performance after engine failure requirements to execute a go-around before LDP or land after LDP but which could not be executed with OEI following an en route engine failure. If such alternate AEO procedures are provided, the Flight Manual should include the appropriate limitations prohibiting use of the AEO procedures after an en route engine failure. For such alternate AEO approach procedures it should be possible to execute a go-around and use the OEI approach procedure if the landing weight is consistent with such approach (the Flight Manual should indicate this OEI approach procedure and corresponding landing weight).

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(ii) Landing distance. Approach and landing path requirements are stated in §§ 29.79(a)(2) and 29.83(a)(2). For Category A rotorcraft, the approach path should allow smooth transition for one-engine inoperative landing and for balked landing maneuvers. For all rotorcraft, the approach and landing paths should allow adequate clearance from potentially hazardous HV combinations. Landing distance is measured from a point 50 feet above the landing surface to a stop. For RFM presentation, the distance is from the aft most portion of the rotorcraft at the point at which the lowest part of the rotorcraft first reaches 50 feet to the foremost point of the rotorcraft (including rotor tip path) after coming to a stop.

(iii) All Engine Out Landing. § 29.79(b) contains the Category A certification requirement for an all engine inoperative landing. The rule states that it should be possible to make a safe landing on a prepared surface after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously.

See paragraph AC 29.143a(2)(iii)(A) for the Category A sequential engine failure criteria. The conditions for last engine failure are maximum continuous power or 30-minute power if that rating is approved, “wings” level flight, and sudden engine failure with a pilot delay of 1 second or normal pilot recognition time, whichever is greater.

Complete power failure has occurred in twin engine Category A rotorcraft. This requirement ensures 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 effected. 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. See discussion under paragraph AC 29.143. The all engine out landing is ordinarily conducted in conjunction with an HV or Category A landing distance phase where ground instrumentation and safety equipment are available. The rotorcraft should be capable of conducting the all engine out landing at the takeoff and landing WAT limiting conditions up to the maximum altitude approved for takeoff and landing.

(2) Procedures.

(i) Instrumentation/Equipment. Instrumentation requirements are basically the same as those for Category A takeoff. A photo theodolite, grid camera, GPS, or other position measuring equipment is needed, along with a ground station to measure wind, OAT, and humidity (if applicable). A two- way communication system between the aircraft and the position measuring equipment is essential. Aircraft instrumentation should include engine and flight parameters, control positions, power lever position, landing gear loads, and a method for synchronizing aircraft position when the power is cut with onboard instrumentation. A record of rotor RPM at touchdown is necessary to ensure 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 support of a fire engine is highly desirable.

(ii) Establishing the LDP.

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(A) Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the LDP becomes largely a function of the runway length required for landing. If landing conditions to be scheduled include considerable runway length (on the order of 1,000 feet), the LDP may be defined at a relatively high speed allowing transition to a takeoff safety speed near VY which will allow the maximum amount of weight for compliance with the balked landing climb requirements of § 29.85(b)/§ 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If the runway length is small, LDP will typically be at a lower speed and may be at a higher altitude to allow balked landing transition within the available distance. Landing weight may need to be reduced to allow landing from the lower speed or higher altitude decision point for shorter landing distances. In this case the requirements of § 29.67(a)(1) may be limiting. The climb performance and climb speeds required by § 29.67(a)(1) and (2) should be established prior to Category A landing tests.

(B) The one-engine-inoperative landing is similar in many respects to the height-velocity tests described in paragraph AC 29.79. Most of the comments, cautio ns, and techniques for HV also apply here even though typical flight conditions at LDP 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 and one engine is made inoperative prior to LDP. After the LDP, speed is reduced and the rotorcraft is flared to a conventional one engine inoperative landing. Depending on the landing characteristics and landing profile, the flare may be initiated either prior or subsequent to the 50 foot elevation utilized in determining landing distance. Testing should include an engine failure such that recognition is at the LDP with a 1-second pilot delay to ensure safe landing capability for this critical case. A sufficient number of acceptable runs should be accomplished to provide confidence in the results. Typically ten acceptable runs are adequate.

(C) The balked landing portion of the landing profile is addressed under § 29.85, Balked Landing: Category A. For an explanation of that requirement and a discussion of those test procedures, refer to paragraph AC 29.77.

(iii) Power. Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping to minimum specification values including consideration of temperature effects on engine power. If the management of engine power at topping is beyond normal pilot capability, the validity of the Cat A procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion the OEI landing, it is acceptable for the engine and transmission transient range in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all Page B - 83

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AC 29-2C, Chg 2 4/25/06 conditions requested by the applicant. Any excursion beyond established transient limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.

(iv) Aircraft Loading. Aft center of gravity is usually most critical for landing distance determination because visibility constraints limit the degree to which the pilot can flare the rotorcraft for landing. 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.

(v) Extrapolation. Landing data may be extrapolated along an established W/σ line to the maximum gross weight of the rotorcraft. However, extrapolation will not be considered valid if landing gear loads are marginally acceptable at actual landing weights below the W/σ limit. If no marginal areas are apparent and an acceptable analytical method is used, performance data may be extrapolated up to 4,000 feet density altitude from test conditions. (See paragraph AC 29.45.)

(vi) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data should be corrected for existing wind conditions during landing distance tests. Credit for headwind conditions may be given during flight manual data expansion.

Paragraph AC 29.1587 details allowable wind credit.

(vii) All Engine Out Landing.

(A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion o f this paragraph, §§ 29.79 and 29.83 each require that a landing from autorotation be possible. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. All -engine-out landing tests should be initiated at light weight with a gradual buildup to the limiting weight conditions. If a complete company test program has documented all -engine-out landings to the GW/σ limit, the buildup conditions during verification test may be decreased. 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.

(B) 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 Page B - 84

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rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which may lead to rotorcraft damage in testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines cut off.

(C) 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. Considerations for weight and altitude extrapolation are the same as those for HV testing (see paragraph AC 29.79). 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.

(D) 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 should stay with allowable limits. Aft center of gravity is ordinarily critical due to visibility and flare- ability.

Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it should be within allowable structural limits.

(viii) Vertical Landings. The reader should be familiar with the preceding discussion of conventional Category A, landing profiles because duplicate information is not repeated here. A typical vertical landing profile is shown in figure AC 29.75A-2.

This profile is equally applicable to both ground level and elevated heliport sites. The profile begins at a stabilized single engine approach condition. It should be possible to make a safe OEI landing or go-around at any point prior to the LDP unless alternate Page B - 85

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AEO approach procedures are presented in the Flight Manual according to paragraph AC 29.75b(1)(i)(A). It is possible to have two landing techniques: an “offset” one, which schedules drop down for elevated heliports (but still ensure 15 feet radial deck edge clearance), and a “straight in” approach which utilizes the ground level heliport criteria. These techniques should be stipulated as such in the Flight Manual. At the LDP the aircraft becomes committed to landing. A safe landing should be possible in case of an engine failure at any point before or after the LDP. Testing should include a simulated failure at LDP with a 1-second delay or normal pilot response time, whichever is longer, and subsequent landing within the allowable area. The landing distance is the distance from the point at which the lowest portion of the rotorcraft reaches 50 feet above the landing surface to the forward-most point after coming to a stop (including main rotor tip path). The LDP becomes very important for landing on small, elevated heliports. The LDP should be clearly defined and Flight Manual instructions should carefully explain any pilot procedures. An illustration similar to figure AC 29.75A-2 with somewhat more detailed information is most useful. Night OEI landings should be conducted to verify suitable visibility for both internal and external vertical landing cues. The minimum elevated heliport size demonstrated for the OEI approach procedure and for alternate AEO approach procedures (when provided) should also be provided in the Flight Manual.

c. Category B Requirements.

(1) Explanation. Section 29.83 contains the Category B landing requirements.

Landing distance is measured from the 50-foot point to the point at which the rotorcraft is completely stopped (approximately 3 knots for water landings). The approach speed is selected by the applicant. Appropriate ambient conditions and allowable extrapolation are discussed under paragraph AC 29.45.

(2) Procedures.

(i) Landing Distance. Aft center of gravity is ordinarily critical due to field-of-view and flare ability. For wheeled rotorcraft, the brakes are applied to an incipient skid for most efficient stopping. For rotorcraft on skids, the collective should be lowered as soon as characteristics allow in order to place a greater weight on the landing skids. These procedures would be appropriate flight manual entries to show how landing distances can be realized. For flight manual purposes, the landing distance should include the horizontal distance from the point at which the lowest part of the rotorcraft first reaches 50 feet above the landing surface to the point at the foremost part of the rotorcraft (including rotor tip path) after coming to a stop. Multiengine rotorcraft incorporating Category A engine isolation features may elect to show compliance with § 29.79 and § 29.81. A sufficient number of acceptable runs should be accomplished to provide confidence in the results. Typically ten acceptable runs are adequate. If a weight effect on landing distance is to be shown, a minimum of two weight extremes are normally tested.

(ii) All-Engine-Out Landing.

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(A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. Section 29.83(c) requires that a landing from autorotation be possible. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface.

All-engine-out landing tests should be initiated at light weight with a gradual buildup to the limiting weight conditions. If a complete company test program has documented all-engine-out landings to the GW/σ limit, the buildup conditions during verification test may be decreased. 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.

(B) 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 low rotor inertia, 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.

(C) 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. Considerations for weight and altitude extrapolation are the same as those for HV testing (see paragraph AC 29.79). 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.

(D) 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 Page B - 87

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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 as 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 should stay with allowable li mits. Aft center of gravity is ordinarily critical due to visibility and flareability. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it should be within allowable structural limits.

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LANDING DISTANCE _____.

FIGURE AC 29.75A-1 CATEGORY A CONVENTIONAL LANDING-CLEAR HELIPORT Page B - 89

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FIGURE AC 29.75A-2 CATEGORY A VERTICAL LANDING Page B - 90

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AC 29.77. § 29.77 (Amendment 29-24) BALKED LANDING: CATEGORY A (For § 29.77 after Amendment 38, see paragraph AC 29.75A)

a. Explanation. This ru le has two distinct portions.

(1) Section 29.77(a) states that the rotorcraft must be capable of transitioning smoothly from each approved Category A approach condition to a missed approach with one engine inoperative (OEI). Although not specifically stated in the rule, this requirement must be met for any point prior to the landing decision point (LDP).

(2) Section 29.77(b) requires that the LDP be defined so that it will permit transition to a safe climb condition in the event a balked landing is necessary. (See figure AC 29.75-1.) The safe climb conditions are defined in § 29.67(a)(1) and (2). This suggests establishing a clearly defined balked landing profile similar to the Category A takeoff profile established under § 29.59. The balked landing profile must insure compliance with the climb performance requirements of §§ 29.67(a)(1) and 29.67(a)(2).

b. Procedures.

(1) Instrumentation. Instrumentatio n requirements are similar to those for Category A takeoff. A ground station with positioning capability is needed along with on-board instrumentation of engine and flight parameters.

(2) Balked Landing Profiles. One engine inoperative balked landing profiles during approach must be conducted at conditions up to and including the LDP. The LDP should be designated so that the balked landing profile may be completed with the rotorcraft descending no lower than 35 feet above the landing surface. The distance from the LDP to the point in the balked landing profile at which a minimum of 35 feet above the landing surface is attained at VTOSS in a climbing posture should be recorded.

This distance should be compared against the landing distance determined under § 29.81 to assure the balked landing maneuver can be completed within the designated landing area. This is especially important for future steep angle, low speed Category A approaches to heliports.

(3) Handling Qualities. Handling qu alities features in the balked landing transition should be carefully evaluated. Characteristics such as excessive nose down pitching with power application or excessive engine lag should not be approved.

(4) Climb Performance. In accordance with this rule, the climb requirements of § 29.67(a)(1) and (2) must also be met in the event a balked landing is made. See paragraphs AC 29.65 and AC 29.67.

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AC 29-2C, Chg 4 5/1/2014 AC 29.79. § 29.79 (Amendment 29-21) LIMITING HEIGHT-SPEED ENVELOPE.

(For § 29.79 after Amendment 29-38, see section 29.75A of this AC.)

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 29.79-1). T he 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) Engine power considerations are similar to those in previous takeoff and landing requirements (see sections 29.53, 29.63, and 29.75 of this AC).

(3) The prohibited sections 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 § 29.143(d)). Altit udes 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 limitation) and normal pilot reaction time.

(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 t hat the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points are approached from above and below respectively. Portions near the knee are initially evaluated at high speed with subsequent backing down of the speed. In most rotorcraft the effective flare airspeed is critical. At airs peeds 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 exerci sed 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.

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(7) For FAA/AUTHORITY tests, the minimum required crew and minimum instrument panel display 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 should be c onsistent with the takeoff and landing weight, altitude, temperature (WAT) limit curve which will be placed in the rotorcraft flight manual (RFM). 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/ (see Figure AC 29.79-2, 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 29.79-2, Part B). Another me thod of presentation is to show varying weights at a constant density altitude (see Figure AC 29.79-2, 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/.

(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.

(11) Rotorcraft certificated prior to Amendment 29-21 were required to have the resulting HV diagram as an operating limitation. This limitation restricted opportunities when operating large rotorcraft in various utility applications. Subsequently, Amendment 29-21 allows, under certain conditions, the HV diagram to be placed in the Flight Manual Performance Information Section instead of the Limitations Section.

Specifically, the rotorcraft must be: (1) certificated for a maximum gross weight of 20,000 pounds or less; (2) configured with nine passenger seats or less; and

(3) certificated in Category B. Testing must be completed with the aircraft at the maximum gross weight at sea level. For altitudes above sea level, the test aircraft must be at a weight no less than the highest weight the rotorcraft can hover out of ground effect (OGE). Rotorcraft certificat ed prior to Amendment 29-21 can update their certification basis to take advantage of this provision.

b. Procedures.

(1) Instrumentation.

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AC 29-2C, Chg 4 5/1/2014

(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, 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 photo panels 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 di agram 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 manufacturer 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 limit. For cruise or “fly-in” points above the knee, the appropriate condition is power required for level flight. Rotor speed at execution of the engine failure should be the minimum speed appropriate to the flight condition.

(ii) The applicable power failure conditions are listed in § 29.79(b). Power should be completely cut for category B 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 adj ustment as ambient conditions change. The power can be takeo ff power (TOP), 2 1/2-minute power, or Page B - 94

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some calculated lower power for simulating hot day or higher density altitude conditions.

Power is verified and recorded by the pilot by “topping” the engine(s) prior to engine failure tests. Care must be taken to assure 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 extrapolatio n is not permitted for HV. Therefore, the test weight must be closely controlled. Ba llast 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 required 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 del ay, 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 hov er 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 mu st 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.

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AC 29-2C, Chg 4 5/1/2014

(ii) Low Hover. From the low hover position there is no flare capability and little time for collective reaction. No ti me 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 very high, unconservative low hover altitudes that are 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 pilo t 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. C onsiderable 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 poin ts 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 high speed touchdown. Test points with excessive gear loads, above average 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 s hould 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 - 96

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response, etc. The HV curve should be pr esented 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. Engine failu res 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. See related discussion under AC 29-2C, section 29.63.

(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.

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HEIGHT (H) ✓-KNEE ~-TAKEOFF

**CORRIDOR**

**HIGH SPEED**

SPEED (V) AC 29-2C, Chg 4 5/1/2014 FIGURE AC 29.79-1 HEIGHT- VELOCITY (HV) DIAGRAM Page B - 98

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Height Velocity Weight

CONSTANT HV DIAGRAM, VARIABLE WEIGHT Height Height Increasing Density Altitude Velocity L_B __J

CONSTANT WEIGHT Increasing Gross Weight Velocity L_c_J Approx.

Constant GW/rr

CON ST ANT DENSITY ALTITUDE FIGURE AC 29.79-2 ALTITUDENVEIGHT ACCOUNTABILITY 5/1/2014 AC 29-2C, Chg 4 Page B - 99

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AC 29.85. § 29.85 (Amendment 29-39) BALKED LANDING: CATEGORY A (For Balked Landing prior to Amendment 39, see § 29.77 and paragraph AC 29.77.)

a. Explanation. Amendment 29-39 revised and relocated the original § 29.77 as a new § 29.85. The guidance material of paragraph AC 29.77 does not apply to rotorcraft certified with Amendment 29-39 or later. This rule has two distinct portions.

(1) Section 29.85(a) states that the rotorcraft must be capable of transitioning smoothly from each approved Category A approach condition to a missed approach with one engine inoperative (OEI). Although not specifically stated in the rule, this requirement must be met for any point prior to the landing decision point (LDP).

(2) Section 29.85(b) requires that the LDP be defined so that it will permit transition to a safe climb condition in the event a balked landing is necessary. (See figure AC 29.75A-1.) The safe climb conditi ons are defined in § 29.67(a)(1) and (2). A clearly defined balked landing profile similar to the Category A takeoff profile should be established. The balked landing profile must insure compliance with the climb performance requirements of §§ 29.67(a)(1) and 29.67(a)(2).

b. Procedures.

(1) Instrumentation. Instrumentatio n requirements are similar to those for Category A takeoff. A ground station with positioning capability is needed along with on-board instrumentation of engine and flight parameters.

(2) Balked Landing Profiles. One engine inoperative balked landing profiles during approach must be conducted at conditions down to and including the LDP. The LDP should be designated so that the balked landing profile may be completed with the rotorcraft clearing the landing surface by a minimum of 15 feet. Fifteen feet should be considered the absolute minimum clearance allowed with greater clearances required for some rotorcraft dependent on rotorcraft geometry and performance characteristics.

For elevated or ground level heliports, with significantly lower LDP heights than 100 feet, the minimum clearance is 15 feet vertically and radially. These minimum heights would need to be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before/after LDP. The distance from the LDP to the point in the balked landing profile at which a minimum of 35 feet above the landing surface is attained at VTOSS in a climbing posture should be recorded. This distance should be com pared against the landing distance determined under § 29.81 to assure the balked landing maneuver can be completed within the designated landing area. This is especially important for future steep angle, low speed Category A approaches to heliports.

(3) Handling Qualities. Handling qu alities features in the balked landing transition should be carefully evaluated. Characteristics such as excessive nose down pitching with power application or excessive engine lag should not be approved.

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(4) Climb Performance. In accordance with this rule, the climb requirements of § 29.67(a)(1) and (2) must also be met in the event a balked landing is made. See paragraphs AC 29.65 and AC 29.67.

AC 29.87. § 29.87 (Amendment 29-39) LIMITING HEIGHT-SPEED ENVELOPE.

(For Limiting Height-Speed Envelope prior to Amendment 39, see § 29.79 and paragraph AC 29.79.)

a. Explanation. Amendment 39 redesignated § 29.79 as § 29.87.

b. Procedures. The guidance material presented in paragraph AC 29.79 continues to apply.

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AC 29-2C, Chg 4 5/1/2014

**SUBPART B - FLIGHT**

FLIGHT CHARACTERISTICS AC 29.141. § 29.141 (Amendment 29-24) FLIGHT CHARACTERISTICS ­ GENERAL.

a. Explanation.

(1) This regulation prescribes the general flight characteristics required for certification of a transport 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 § 29.141(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 V Y, 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 §§ 29.141 through 29.161 are met at takeoff power levels, no flight manual entry is needed. Obviously, if transmission limits for maximum continuous (MC) and takeoff power are coincident, 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 § 29.143(c) would be in question. The same consideration must be given to all rotorcraft controls and moveable 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 § 29.251.

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(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 29.695 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 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, it should be considered if this failure mode should also be investigated during high altitude testing 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 29.141(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, § 29.141(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.

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AC 29-2C, Chg 4 5/1/2014 AC 29.143. § 29.143 (Amendment 29-24) CONTROLLABILITY AND MANEUVERABILITY.

a. Explanation.

(1) This regulation contains the basic controllability requirements for transport rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for control and for maneuverability are summarized in § 29.143(a) which is largely self-explanatory. During the assessment carried out under § 29.143(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 § 29.143(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), § 29.143, include more specific flight conditions and highlight the typical areas of concern during a flight test program.

(i) Section 29.143(b) specifies flight at VNE with critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at V NE 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 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 29.143(c) requires a minimum 17-knot control capability for hover and takeoff in winds 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 at all altitudes approved for takeoff and landing. On rotorcraft incorporating a tail rotor, Page B - 104

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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. Prior to Amendment 29-24, hover controllability, height-velocity, and hover performance were the three regulatory requirements that ordinarily determined the shape of the limiting weight-altitudetemperature (WAT) curve for takeoff and landing. For category A performance rotorcraft operations, of course, the one engine inoperative (OEI) climb performance requirements may also influence the WAT limit curve. Amendment 29-24 allows, under certain conditions, the deletion of any hover controllability condition determined under § 29.143(c) from becoming an operating limitation. Section 29.1587 of Amendment 29- 24 provides a means wherein category B certificated rotorcraft (in accordance with the requirements of § 29.1, effective with Amendment 29-21) may not be limited by the hover controllability requirements of § 29.143(c). Section 29.1583(g) requirements for category A certificated rotorcraft are unchanged from past regulatory requirements in that if the hover controllability requirements of § 29.143(c) result in the most restrictive envelope it will be published as an operating limitation. Section 29.1587(b) provides a means wherein category B certificated rotorcraft, as defined in § 29.1, may not be restricted in its utilization. It allows such rotorcraft to publish the maximum takeoff and landing capabilities of the rotorcraft, provided something other than the 17 knot hover controllability requirement is not limiting. This may be zero wind IGE hover performance or any other performance the applicant elects to use if the maximum safe wind for operations near the ground is provided. Rotorcraft certificated prior to Amendment 29- 24 can update their certification basis to take advantage of this provision. If an applicant with a previously type certificated rotorcraft elects to update to this later amendment, caution should be taken to verify that the HV information is done in accordance with Amendment 29-21; that all engine out landing capabilities are satisfactorily accounted for at the new proposed gross weight, altitude, temperature combinations; that takeoff and landing information is provided; and that sufficient information is provided to properly advise the crew of the rotorcraft’s capabilities when utilizing this increased performance capabilities.

(iii) Section 29.143(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 which meet the engine isolation requirements of 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 c onditions for the first engine failure are engine or transmission limiting maximum continuous power (MCP) (or take-off 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 § 29.75(b)(5).

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AC 29-2C, Chg 4 5/1/2014 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 category B powerplant installation rotorcraft, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to V NE (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 part 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 in order to comply with § 29.63(c). Following power failure, rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.

(iv) Section 29.143(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 § 29.1505(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 V NE at MCP per § 29.143(b), 1.1 VNE at power for 0.9 VH per § 29.175(b) or § 29.1505, and to 1.1 VNE (power-off) in autorotation per § 29.143(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 Page B – 104.2

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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 (§ 29.672). 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 they 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 RFM.

b. Procedures.

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AC 29-2C, Chg 4 5/1/2014

(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 1.1 VH) point obtained during stability tests required by § 29.175(b).

Controllability must be satisfactory for both conditions. If V NE 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 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 takeoff power (TOP) 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 Page B – 104.4

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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 they should be specified in the TIA. Particular attention should be given to minimum stabilized and minimum transient rotor RPM limits. These values must be included in the TIA and should be approached gradually with a build-up in time delay unless the company testing has completely validated all pertinent aspects of engine failure testing. On category A installations the maximum power output of each engine must be limited so that when an engine fails and the remaining engine(s) assume the additional load, the remaining engine(s) are not damaged by excessive power extraction and over-temping. This is needed for compliance with § 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 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 Page B – 105

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AC 29-2C, Chg 4 9/17/2009 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 § 29.75 (see section 29.75 of this AC).

AC 29.143A. § 29.143 (Amendment 29- 51) Controllability and Maneuverability.

a. Explanation. Amendment 29- 51 made a minor clarification to assure that inground-effect (IGE) controllability is demonstrated at all wind speeds up to 17 knots, for all azimuths. 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 out -of-ground-effect (OGE) controllability be determined up to a speed of at least 17 knots at a weight selected by the applicant. The amendment clarifies the intent of Amendment 29- 21 and Amendment 29-24 with respect to removing hover controllability as a limit.

Section 29.25 is amended to assure that appropriate weight limitations are incorporated into the Rotorcraft Flight Manual (RFM) when the relieving provisions of the previous amendments are adopted by an applicant. The previous amendment and associated AC material indicated that certain Category B rotorcraft were relieved from providing, as a limitation, the conditions of § 29.143(c). In practice, the 17- knot controllability requirement was still treated as a limitation, but, as indicated in the amended § 29.25, additional limits could be included, when demonstrated, that allowed for something other than 17-knot all azimuth controllability. The established weight, altitude, and temperature charts, including any associated wind constraints, could be contained in the performance section of the flight manual when the appropriate reference to those charts were included in the limitations section of the RFM. In addition, the relief of Amendments 29-21 and 29-24 were only intended for those category B rotorcraft with nine or less passenger seats. All the policy material pertaining to this section remains in effect with the following changes:

(1) This regulation contains the basic controllability requirements for transport rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for controllability and for maneuverability are summarized in § 29.143(a) which is self-explanatory. The hover condition is not specifically addressed in § 29.143(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 § 29.143 include more specific flight conditions and highlight the typical areas of concern during a flight test program.

(i) Section 29.143(b) specifies flight at VNE with critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at V NE for Page B – 106

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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 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) Section 29.143(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 near the ground such as rolling takeoffs for wheeled rotorcraft. These requirements must be met at all altitudes approved for takeoff and landing. 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. Prior to Amendment 29-24, hover controllability, height-velocity, and hover performance were the three regulatory requirements that ordinarily determined the shape of the limiting weight-altitude-temperature (WAT) curve for takeoff and landing. For Category A performance rotorcraft operations, of course, the one-engine-inoperative (OEI) climb performance requirements may also influence the WAT limit curve. Amendment 29- 24 allows, under certain conditions, the deletion of any hover controllability condition determined under § 29.143(c) from becoming an operating limitation. Section 29.1587 of Amendment 29-24 provides a means wherein Category B certificated rotorcraft (in accordance with the requirements of § 29.1, effective with Amendment 29-21) may not be limited by the hover controllability requirements of § 29.143(c). Section 29.1583(g) requirements for Category A certificated rotorcraft are unchanged from past regulatory requirements in that if the hover controllability requirements of § 29.143(c) result in the most restrictive envelope it will be published as an operating limitation. Section

29.1587(b) provides a means wherein Category B certificated rotorcraft, as defined in § 29.1, may not be restricted in its utilization. Section 29.1587(b) allows some Category B RFMs to include maximum takeoff and landing performance information, provided that something other than the 17-knot hover controllability requirement is not limiting. This may be zero wind IGE hover performance or any other performance the applicant elects to use, if the maximum safe wind for operations near the ground is provided. Rotorcraft certificated prior to Amendment 29-24 can update their certification basis to take advantage of this provision. If an applicant with a previously type certificated rotorcraft elects to update to this later amendment, caution should be taken to verify that the Page B – 106.1

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AC 29-2C, Chg 4 9/17/2009 height-velocity information is done in accordance with Amendment 29-21; that all engine out landing capabilities are satisfactorily accounted for at the new proposed gross weight, altitude, temperature combinations; that takeoff/landing information is provided;

and that sufficient information is provided to properly advise the crew of the rotorcraft's capabilities when utilizing this increased performance capabilities.

(iii) Section 29.143(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 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 § 29.75(b)(5). 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 Category B powerplant installation rotorcraft, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to V NE (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 part 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 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 in order to comply with § 29.63(c). Following power failure, the rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.

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(iv) Section 29.143(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 § 29.1505(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 V NE at MC power per § 29.143(b) and § 29.175(c), and to 1.1 VNE (power-off) in autorotation per § 29.143(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 RFM.

(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 29.672). 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 they are subject to actuator softover and hardover Page B – 106.3

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AC 29-2C, Chg 4 9/17/2009 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. A n 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 1.1 VH ) point obtained during stability tests required by § 29.175(b).

Controllability must be satisfactory for both conditions. If V NE 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 – 106.4

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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 rotorcraft 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 rotorcraft 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. There may be other conditions where adequate yaw control is not available at high power, for example a rotorcraft which is limited by the CAT A weight (for rotorcraft certificated to § 29.1 (c)).

(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. This will account 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.

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AC 29-2C, Chg 4 9/17/2009

(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 29.45 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 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 they should be specified in the 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 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 pitch-up is satisfactory. Obviously Page B – 106.6

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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 requirements of § 29.143(a)(2)(v) and (vi).

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AC 29-2C, Chg 4 5/1/2014 AC 29.151. § 29.151 (Amendment 29-24) FLIGHT CONTROLS.

a. Explanation. Excessive breakout or preload in the flight controls produces control system force discontinuities, which result in increased workload and even 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. Although in some designs friction can provide a positive contribution to the function of the flight controls (e.g., masking aerodynamic feedback in reversible systems), friction will eventually 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.

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AC 29.161. § 29.161 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 provide the pilot with a refer ence cyclic control position for the given flight condition, reduce the physical demands to maintain a given flight condition, and allow the pilot to release the cyclic control for brief periods of time to perform other cockpit duties. A primary flight control which can move when released imposes an additional pilot workload by requiring a continuous hands-on 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, simply 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 assured 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 § 29.161. For additional IFR trim requirements, refer to AC 29 Appendix B.

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AC 29-2C, Chg 4 5/1/2014 AC 29.161A. § 29.161 (Amendment 29-24) TRIM CONTROL.

a. Explanation. Amendment 29- 24 to the regulation adds the additional requirement that the trim control be capable of trimming collective forces to zero.

b. Procedures. The trim requirement is intended to allow the pilot to release the controls for brief periods to perform other cockpit duties, and to provide the pilot with a reference cyclic position for the given flight condition. The collective should be balanced so that there is no tendency for the collective pitch to change when the collective is released. Any magnetic clutch, friction brake or similar device which modifies the collective characteristics should be capable of being overpowered by the pilot, when fully applied, without requiring excessive force.

AC 29.171. § 29.171 STABILITY: GENERAL.

a. Explanation. This section is intended to require a manageable pilot workload for the minimum crew under foreseeable operating conditions.

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 flight crew.

(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 (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 assure pilot strain and fatigue are acceptable during typical flight profiles for each type of operation to be approved.

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AC 29.173. § 29.173 (Amendment 29-24) STATIC LONGITUDINAL STABILITY.

a. Explanation.

(1) This 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 which 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.

(2) The remainder of § 29.173, through reference to § 29.175, 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 § 29.175 by requiring a stable stick position gradient through a specified speed range. A defined level of instability is permitted for the hovering condition.

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 assure compliance with the stability requirements of this section are contained under § 29.175, Demonstration of static longitudinal stability. Refer to paragraph AC 29.175 for an explanation of detailed flight test procedures.

AC 29.173A. § 29.173 (Amendment 29- 51) STATIC LONGITUDINAL STABILITY.

a. Explanation.

(1) Amendment 29-51 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 § 29.173(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 Page B – 111

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