# AC 29-2B: Certification of Transport Category Rotorcraft

- Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it)
- Status: Historical
- Version in force: [AC 29-2C CHG 9](https://openairdocs.org/ac/29-2), approved June 23, 2023. This is a historical version
- Approved: July 30, 1997
- Last reviewed: January 23, 2001
- Family: AC 29-2
- Cancels: [AC 29-2A](https://openairdocs.org/ac/29-2A)
- Cancelled by: [AC 29-2C](https://openairdocs.org/ac/29-2C)
- 14 CFR parts: 29
- Pages: 912
- PDF: https://openairdocs.org/files/sha256/d1/a6/d1a61c181e6b7f48427ad1864b56450b6ba10ac8ed3cf4bb660175b46045088d
- FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/9F988435EB93C0FE862569DD00606CDB.0001
- DRS record last changed: November 22, 2025
- Page: https://openairdocs.org/ac/29-2B

## Versions

- [AC 29-2C CHG 9](https://openairdocs.org/ac/29-2) (Current), approved Jun 23, 2023, 1,453 pages
- [AC 29-2C CHG 4-8](https://openairdocs.org/ac/29-2C-CHG-4-8) (Historical), approved Jul 2, 2018, 1,392 pages
- [AC 29-2C CHG 1-3](https://openairdocs.org/ac/29-2C-CHG-1-3) (Historical), approved Sep 30, 2008, 1,146 pages. Changes 1-3
- [AC 29-2C CHG 2](https://openairdocs.org/ac/29-2C-CHG-2) (Historical), approved Apr 25, 2006, 1,139 pages
- [AC 29-2C](https://openairdocs.org/ac/29-2C?v=3B7C96A86441B537862570380068402E) (Historical), approved Jun 24, 2005, 17 pages. CHGS to Various Paragraphs
- [AC 29-2C](https://openairdocs.org/ac/29-2C?v=9AEB86448234C61886256D4200646908) (Historical), approved Apr 25, 2003, 4 pages
- [AC 29-2C](https://openairdocs.org/ac/29-2C?v=2648447DA95F593886256D4200646926) (Historical), approved Apr 25, 2003, 3 pages
- [AC 29-2C](https://openairdocs.org/ac/29-2C?v=0AA79C1893E613F886256D4200646944) (Historical), approved Apr 25, 2003, 4 pages
- [AC 29-2C CHG 1](https://openairdocs.org/ac/29-2C-CHG-1) (Historical), approved Feb 12, 2003, 2 pages
- [AC 29-2C](https://openairdocs.org/ac/29-2C?v=29E440234B4CC27586256E4300597DD7) (Historical), approved Feb 12, 2003, 1,109 pages
- [AC 29-2B](https://openairdocs.org/ac/29-2B) (Historical), approved Jul 30, 1997, 912 pages
- [AC 29-2 CHG 1](https://openairdocs.org/ac/29-2-CHG-1) (Historical), approved Nov 19, 1984, 118 pages
- [AC 29-2C](https://openairdocs.org/ac/29-2C?v=405A2C0BF7BF68C58625707B00718417) (Historical), approved , 4 pages. 29.351B - Yawing Conditions

## Text

Each page's text from the PDF; open the PDF at a page with #page=N.

### Page 1

AdvisoryU.S.Department ofTmsportation F8deml Aviation Mminlstmtion Circular Subject: CERTIFICATIONOF TRANSPORT Date: 7/30/97 AC No: 29-2B CATEGORY ROTORCRAFT Initiatedby: ASW-110 Change:

1. PURPOSE:

a. This is a total revision of AC 29-2A dated 9/16/87, with changes 1, 2, and 3, dated 4/24/89, 9/24/91, and 6/1/95 respectively, incorporated. In addition, new material plus changes to existing paragraphs have been incorporated. This consolidated version is now renumbered as AC 29-2B and replaces AC 29-2A in its entirety. This revises existing material in 25 paragraphs and adds new material for 33 paragraphs.

b. 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.

c. 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 29-2A, Certification of Transport Category Rotorcraft, September 16, 1987, is canceled in its entirety.

3. BACKGROUND. Based largely on precedents set during rotorcraft certification programs spanning the past 39 years, this AC consolidates guidance contained in earlier correspondence among FAA headquarters, foreign authorities, the rotorcraft industry, and certificating regions.

4. PRINCIPAL CHANGES:

a. Paragraphs 31A} 32, 45, 47, 55, 57, 64, 69, 71, 72, 140A, 245, 337, 596, 618, 619, 621, 633, 641, 652, 653, 726, 765, 775, and 777 are revised to incorporate technical guidance.

b. New paragraphs 42A, 55B, 56, 57A, 58A, 59, 60A, 66A, 67A, 70A, 71A, 72A, 140B, 152A, 205A, 218B, 252A, 254, 329B, 359A, 397B, 398C, 421A, 423C, 447, 454B, 456A, 459A, 460B, 563B, 619B, 619C, 724B, and 765A are added to Chapter 2.

c. New paragraph 781A is added to Chapter 3.

d. Paragraph 447, ~ 29.951, General, is renumbered to Paragraph 446. Paragraph 447 now addresses ~ 29.952, Fuel Systems Crash Resistance.

### Page 2

e. The following appendices have been added:

Appendix 2 One-Engine-Inoperative (OEI) Power Assurance Appendix 3 Rotorburst

f. Use of the term “FAA/AUTHORITY” replaces “FAA as appropriate. “FAA/AUTHORITY” as used in this document means FAA or another ainvorthiness authority that has adopted this AC as a means of compliance with the appropriate regulation referenced.

5. DEVIATIONS. As rotorcraft designs vary from conventional configurations, it may become necessary to deviate from the methods and procedures outlined in this AC. These procedures are only one acceptable means of compliance with Part 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-I 10, or the appropriate ai~orthiness 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 Aircraft Certification Service

### Page 3

**TABLE OF CONTENTS CHAPTER 1- PART 21 CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS**

1.-3.

4.

5.-7 &

9.-11.

12.

13.-15.

16.

17.-23.

24.

25.-30.

Page No.

**RESERVED**

FAR 21.16 Special Conditions … 5

**RESERVED**

FAR21 .31 Type Design … 9

**RESERVED**

FAR21 .33 Inspection and Tests … l3

**RESERVED**

FAR21.35 Flight Tests … l7

**RESERVED**

FAR 21.39 Flight Test Instrument Calibration and … 27

Correction Report

RESERVED CHAPTER 2- PART 29 AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY

**ROTORCRAFT SECTION 1. GENERAL**

31. FAR 29.1 Applicability … 35

31A. FAR 29.1 Applicability … 36

32. FAR 29.2 Special Retroactive Requirements … 37

**33.-41. RESERVED WTION 2, FLIGHT - GENERAL**

42. FAR29.21 Proof of Complian= … 48

42A. FAR29.21 Proof of Compliance … 50

43. FAR29.25 Weight Limits … 50

44. FAR29.27 Center of Gravity Limits … 52

45. FAR 29.29 Empty Weight and Corresponding … 53

Center of Gravity

46. FAR29.31 Removable Ballast … 54

47. FAR29.33 Main Rotor Speed and Pitch Limits … 55

48.-54. RESERVED Ill

### Page 4

**SECTION 3. PERFORMANCE**

55.

55A.

55B.

56.

57.

57A.

58.

58A.

59.

60.

60A.

Page No.

FAR 29.45 General … 64

FAR 29.45 General … 74

FAR 29.45 Generai … 74

FAR 29.49 Performance at Minimum Operating Speed. Hover … 81

Performance for Rotorcraft

FAR 29.51 Takeoff Data: General … 83

FAR 29.51 Takeoff Data: General … 84

FAR29.53 Takeoff Categoy A … 85

FAR29.53 Takeoff Categoy A … 88

FAR29.55 Takeoff Decision Point: Catego~A … 9O

FAR 29.59 Takeoff Path: Category A … 90

FAR 29.59, 29.60, 29.61, and29.62 Takeoff Path, Distance ...................................l Ol and Rejected Takeoff Ground Level and Elevated Heliport:

Category A

**61.-63. RESERVED**

64.

65.

66.

66A.

67.

67A.

68.

69.

70.

70A.

71.

71A.

72.

72A.

FAR 29.63 Takeoff Path: Category B … 115

**RESERVED**

FAR29.65 Climb: All Engines Operating … ll9

FAR 29.64 and 29.65 Climb: (General and All Engines operating) … 121

FAR29.67 Climb: One Engine inoperative … l23

FAR 29.67 Climb: One Engine inoperative … l24

FAR 29.71 Rotorcraft Angle of Glide: Category B … l26

FAR 29.73 Performance at Minimum Operating Speed … 129

Hover Performance for Rotorcraft

FAR29.75Landing … l3l

FAR 29.75, 29.77, 29.79, 29.81, and 29.83 Landing … l4l

FAR 29.77 Balked Landing: Catego~A … l52

FAR 29.85 Balked Landing: Catego~A … l53

FAR 29.79 Limiting Height-Speed Enveiope … l54

FAR 29.87 Limiting Height-Speed Envelope … l62

**73.-78. RESERVED - - SECTION 4, FLIGHT CHARACTERISTICS**

79. FAR 29.141 Flight Characteristics .General … l69

80. FAR 29.143 Controllability and Maneuverability … l7O

81. FAR 29.151 Flight Controls … l76

82. FAR 29.161 Trim Control … l76

82A. FAR 29.161 Trim Control … l77

83. FAR 29.171 Stability: General … l78

84. FAR29.173 Static Longitudinal Stability … l78

Iv

### Page 5

SECTION 4. FLIGHT CHARACTERISTICS (continued) Page No.

85. FAR 29.175 Demonstration of Static Longitudinal Stability … 179

86. FAR 29.177 Static Directional Stability … 184

87. FAR 29.181 Dynamic Stability: Category A Rotorcrafi … l85

**88.-95. RESERVED SECTION 5. G ROUND AND WATER HANDLING CHARACTERISTICS**

96. FAR 29.231 General (Ground and Water Handling … l94

Characteristics)

97. FAR29.235 Taxiing Condition … l94

98. FAR 29.239 Spray Characteristics … l95

99. FAR 29.241 Ground Resonance … l96

**100.-109. RESERVED SECTION 6. MISCELLANEOUS FLIGHT REQUIREMENTS**

110. FAR 29.251 Vibration … 2l2

**111.-120. RESERVED SECTION 7. STRENGTH REQUIREMENTS - GENERAL**

121. FAR 29.301 Loads … 225

122. FAR29.303Factor of Safety … 225

123. FAR 29.305 Strength and Deformation … 226

124. FAR29.307Proof of Structure … 226

124A. FAR 29.307 Proof of Structure … 227

125. FAR 29.309 Design Limitations … 228

**126.-135. RESERVED SECTION 8. FLIGHT LOADS**

136. FAR 29.321 General-Flight Loads … 239

137. FAR 29.337 Limit Maneuvering Load Factor … 239

138. FAR 29.339 Resultant Limit Maneuvering Loads … 242

139. FAR 29.341 Gust Loads … 242

140. FAR29.351 Yawing Conditions … 244

140A. FAR 29.351 Yawing Conditions … 247

140B. FAR 29.351 Yawing Conditions … 249

141. FAR 29.361 Engine Torque … 249

142.-151. RESERVED v

### Page 6

SECTIO N9. CONTROL SURFACE AND SYSTEMS LOADS Page No.

152. FAR29.391 General … 26l

152A. FAR 29.391 General … 261

153. FAR 29.395 Control System … 26l

153A. FAR 29.395 Control System … 263

154. FAR 29.397 Limit Pilot Forces and Torques … 263

155. FAR 29.399 Dual Control System … 264

156. FAR 29.401 Auxiliary Rotor Assemblies … 264

156A. FAR 29.401 Auxiliary Rotor Assemblies … 265

157. FAR 29.403 Auxiliary Rotor Attachment Structure … 266

157A. FAR 29.403 Auxiliary Rotor Attachment Structure … 266

158. FAR 29.411 Ground Clearance: Tail Rotor Guard … 267

159. FAR 29.413 Stabilizing and Control Surfaces … 268

159A. FAR 29.413 Stabilizing and Control Surfaces … 270

**160.-161. RESERVED**

162. FAR29.427 Unsymmetrical Loads … 273

**163.-169. RESERVED SEC TION 10. GRO UND LOADS**

170. FAR 29.471 General … 284

171. FAR 29.473 Ground Loading Conditions and Assumptions … 284

172. FAR 29.475 Tires and Shock Absorbers … 285

173. FAR29.477 Landing Gear Arrangement … 285

174. FAR29.479 Level Landing Conditions … 285

175. FAR 29.481 Tail-Down Landing Conditions … 287

176. FAR 29.483 One-Wheel Landing Conditions … 287

177. FAR 29.485 Lateral Drift Landing Conditions … 287

178. FAR 29.493 Braked Roll Conditions … 288

179. FAR 29.497 Ground Loading Conditions: Landing Gear With … 288

Tail Wheels

180. FAR 29.501 Ground Loading Conditions: Landing Gear With … 290

Skids

180A. FAR 29.501 Ground Loading Conditions: Landing Gear With … 292

Skids

181. FAR29.505Ski Landing Conditions … 292

182. FAR 29.511 Ground Load: Unsymmetrical Loads on Multiple … 293

Wheel Units

**183.-192. RESERVED VI**

### Page 7

SECTION 11. WATER LOADS Page No.

193. FAR 29.519 Hull Type Rotorcraft: Water Based, Amphibian … 304

194. FAR 29.521 Float Landing Conditions … 3O6

**195.-204. RESERVED SECTION 12. MAIN COM ONENP T RFQUI REMENTS**

205. FAR 29.547 Main Rotor Structure … 3l7

205A. FAR 29.547 Main Rotor and Tail Rotor Structure … 3l7

206. FAR 29.549 Fuselage and Rotor Pylon … 32l

207. FAR 29.551 Auxiliary Lifting Surfaces … 32l

**208.-217. RESERVED SECTION 13, EMERGE NCY I ANDING CO NDITIONS**

218. FAR29.561 General … 332

218A. FAR 29.561 Emergency Landing Conditions .General … 334

218B. FAR29.561 Emergency Landing Conditions .General … 334

219. FAR 29.562 Emergency Landing Dynamic Conditions … 335

220. FAR 29.563 Structural Ditching Provisions … 335

220A. FAR 29.563 Structural Ditching Provisions … 335

**221.-229. RESERVED SECTION 14. FATIGU E EVALUATION**

230. FAR 29.571 Fatigue Evaluation of Flight Structure … 347

230A. FAR 29.571 Fatigue Evaluation of Flight Structure … 349

**231.-239. RESERVED**

240.

241.

242.

243.

244.

245.

246.

247.

**SECTION 15. DESIGN AND CONSTRUCTION - GENERAL**

FAR29.601 Design … 36O

FAR29.603Materials … 36O

FAR29.605 Fabrication Methods … 36l

FAR 29.607 Fasteners … 362

FAR 29.609 Protection of Structure … 363

FAR 29.610 Lightning and Static Electricity Protection … 364

FAR 29.611 Inspection Provisions … 37O

FAR 29.613 Material Strength Properties and Design … 37l

Values Vll

### Page 8

SFCTION 15. DESIGN AND CONSTRU CTION - GENERAL (continued] Page No.

247A. FAR 29.613 Material Strength Properties and Design … 373

Values

248. FAR 29.613 Special Factors … 373

249. FAR 29,621 Casting Factors … 374

250. FAR 29.623 Bearing Factors … 384

251. FAR 29.625 Fitting Factors … 385

252. FAR 29.629 Flutter … 386

252A. FAR29.629 Flutter and Divergence … 386

**253. RESERVED**

254. FAR29.631 Bird Strike … 38g

**255.-264. RESERVED SECTION 16, ROTORS**

265. FAR 29.653 Pressure Venting and Drainage of Rotor Blades … 400

266. FAR29.659 Mass Balance … 4OO

267. FAR29.661 Rotor Blade Clearance … 4OO

268. FAR29.663 Ground Resonance Prevention Means … 401

268A, FAR29.663 Ground Resonance Prevention Means … 403

**269.-275. RESERVED SECTIO N 17. CO NTROL SYSTEMS**

276.

277.

278.

279.

280.

281.

282.

283.

284.

285.

286.

287.

FAR29.671 General … 4l2

FAR 29,672 Stability Augmentation, Automatic … 412

and Power-Operated Systems

FAR29.673Prima~ Flight Controls … 4l3

FAR29.674 Interconnected Controls … 4l4

FAR29.675Stops … 4l4

FAR29.679 Control System Locks … 4l6

FAR 29.681 Limit Load Static Tests … 4l7

FAR29.683 Operation Tests … 4l7

FAR29.685 Control System Details … 4l8

FAR 29.687 Spring Devices … 42O

FAR 29.691 Autorotation Control Mechanism … 42O

FAR 29.695 Power Boost and Power-Operated Control System … 421

**288.-297. RESERVED**

### Page 9

7/30/97 SECTION 18, LA NDING GEAR AC 29-2B Page No.

298. FAR 29.723 Shock Absorption Tests … 434

299. FAR 29.725 Limit Drop Test … 437

300. FAR 29.727 Reserve Energy Absorption Drop Test … 437

300A. FAR 29.727 Reserve Energy Absorption Drop Test … 438

301. FAR 29.729 Retracting Mechanism, Landing Gear … 439

302. FAR 29.731 Wheels … 44l

303. FAR 29.733 Tires ................0 … 441

304. FAR 29.735 Brakes … 443

305. FAR 29.737 Skis … 445

**306.-315. RESERVED SECT ION 19. FLOATS AND HULLS**

316. FAR 29.751 Main Float Buoyancy … 457

317. FAR 29.753 Main Float Design … 457

318. FAR 29.755 Hull Buoyancy … 46O

319. FAR 29.757 Hull and Auxiliary Float Strength … 462

**320.-326. RESERVED SEC TION 20. PERSONNEL AN D CARGO ACCO MMODATIONS**

327. FAR 29.771 Pilot Compartment … 470

328. FAR29.773Pilot Compadment View … 47l

329. FAR 29.775 Windshields and Wndows … 477

329A. FAR 29.775 Windshields and wndows … 478

329B. FAR 29.775 Windshields and Wndows … 478

330. FAR 29.777 Cockpit Controls … 478

331. FAR 29.779 Motion and Effect of Cockpit Controls … 48O

**332.-333. RESERVED**

334. FAR 29.783 Doors … 485

334A. FAR 29.783 Doors … 488

335. FAR 29.785 Seats, Safety Belts, and Harnesses … 489

335A. FAR 29.785 Seats, Berths, Belts, Safety Belts, and … 49l

Harnesses

336. FAR 29.787 Cargo and Baggage Compadments … 493

336A. FAR 29.787 Cargo and Baggage Compartments .......................................o … 494

337. FAR 29.801 Ditching … 494

lx

### Page 10

SECTION 20. PERSONNEL AND CARGO ACCOMMODATIONS (continued) Page No.

338. FAR 29.803 Emergency Evacuation … 5ol

338A. FAR 29.803 Emergency Evacuation … 5O2

339. FAR 29.805 Flightcrew Emergency Exits … 5o3

339A. FAR 29.805 Flightcrew Emergency Exits … 5o3

340. FAR 29.807 Passenger Emergency Exits … 5o4

340A. FAR 29.805 Emergency Exits … 505

341. FAR 29.809 Emergency Exit Arrangement … 506

341A. FAR 29.809 Emergency Exit Arrangement … 5o7

342. FAR 29.811 Emergency Exit Marking … 5O8

342A. FAR 29.811 Emergency Exit Marking … 510

343. FAR 29.812 Emergency Lighting … 511

344. FAR 29.813 Emergency Exit Access … 511

345. FAR 29.815 Main Aisle Width … 5l3

346. FAR 29.831 Ventilation … 5l4

347. FAR 29.833 Heaters … 515

**348.-356. RESERVED SECTION 21. FIRE PROTECTION**

357. FAR29.851 Fire Extinguishers … 526

358. FAR29.853 Compartment interiors … 526

359. FAR29.855Cargo and Baggage Compatiments … 529

359A. FAR 29.855 Cargo and Baggage Compafiments … 53l

360. FAR 29.859 Combustion Heater Fire Protection ..................................................53 I

361. FAR 29.861 Fire Protection of Structure, Controls, and … 535

Other Parts

361A. FAR 29.861 Fire Protection of Structure, Controls, and … 538

Other Parts

362. FAR 29.863 Flammable Fluid Fire Protection … 538

**363.-372. RESERVED SECTION 22, EXTERNAL LOAD ATTACHING MEANS**

373. FAR 29.865 External Load Attaching Means … 55l

373A. FAR 29.865 External Load Attaching Means … 556

374.-383. RESERVED x

### Page 11

SECTION 23. MISCE OUS ( )LLANE ,DESIGN AND CONSTRUCTION Page No.

384. FAR 29.871 Leveling Marks … 576

385. FAR 29.873 Ballast Provisions … 576

386. FAR 29.877 Ice Protection … 577

**387.-396. RESERVED SECTIO N 24. POWERPLANT - GENERAL**

397. FAR 29.901 Installation … 608

397A. FAR 29.901 Installation … 615

397B. FAR 29.901 Installation .....................o … 6l5

398. FAR 29.903 Engines … 616

398A. FAR 29.903 Engines … 6l8

398B. FAR 29.903 Engines … 620

398C. FAR 29.903 Turbine Engine Installation … 62O

399. FAR 29.907 Engine Vibration … 62O

400. FAR 29.908 Cooling Fans … 621

400A. FAR 29.908 Cooling Fans … 622

**401.-420. RESERVED SECTION 25. ROTOR 17RIVE SYSTEM**

421. FAR 29.917 Design … 645

421A. FAR 29.917 Design … 646

422. FAR 29.921 Rotor Brake … 649

423. FAR 29.923 Rotor Drive System and Control Mechanism … 65l

Tests

423A. FAR 29.923 Rotor Drive System and Control Mechanism … 657

Tests

423B. FAR 29.923 Rotor Drive System and Control Mechanism … 664

Tests

423C. FAR 29.923 Rotor Drive System and Control Mechanism … 664

Tests

424. FAR 29.927 Additional Tests … 666

424A. FAR 29.927 Additional Tests … 668

425. FAR 29.931 Shafting Critical Speeds … 67l

426. FAR 29.935 Shafting Joints … 676

427. FAR 29.939 Turbine Engine Operating Characteristics … 676

428.-445. RESERVED xl

### Page 12

AC 29-26 SECTION 26. FUEL SYSTEM 7/30/97 Page No.

446. FAR 29.951 General … 697

447. FAR 29.952 Fuel System Crash Resistance … 699

448. FAR 29.953 Fuel System independence … 725

449. FAR 29.954 Fuel System Lightning Protection … 726

450. FAR 29.955 Fuel Flow … 727

450A. FAR 29.955 Fuel Flow … 731

451. FAR29.957 Flow Between Interconnected Tanks … 735

452. FAR 29.959 Unusable Fuel Supply … 735

453. FAR 29.961 Fuel System Hot Weather Operation … 736

453A. FAR 29.961 Fuel System Hot Weather Operation … 738

454. FAR29.963 Fuel Tanks: General … 739

454A. FAR29.963 Fuel Tanks: General … 739

4546. FAR29.963 Fuel Tanks: General … 74O

455. FAR29.965Fuel Tank Tests … 74l

456. FAR 29.967 Fuel Tank Installation … 743

456A. FAR 29.967 Fuel Tank installation … 745

457. FAR 29.969 Fuel Tank Expansion Space … 746

457A. FAR 29.969 Fuel Tank Expansion Space … 746

458. FAR 29.971 Fuel Tank Sump … 746

458A. FAR 29.971 Fuel Tank Sump … 748

459. FAR29.973Fuel Tank Filler Connection … 749

459A. FAR 29.973 Fuel Tank Filler Connection … 749

460. FAR 29.975 Fuel Tank Vents and Carburetor Vapor Vents … 751

460A. FAR 29.975 Fuel Tank Vents and Carburetor Vapor Vents … 751

4606. FAR 29.975 Fuel Tank Vents and Carburetor Vapor Vents … 752

461. FAR29.977Fuel Tank Outlet … 753

462. FAR 29.979 Pressure Refueling and Fueling Provisions Below Fuel Level … 753

**463.-482. RESERVED SECTION 27. FUEL SYSTEM COMPONENTS**

483. FAR29.991 Fuel Pumps … 775

483A. FAR29.991 Fuel Pumps … 776

484. FAR 29.993 Fuel System Lines and Fittings … 776

485. FAR29.995 Fuel Valves … 777

486. FAR 29.997 Fuel Strainer or Filter … 777

486A. FAR 29.997 Fuel Strainer or Filter … 778

487. FAR 29.999 Fuel System Drains … 779

487A. FAR 29.999 Fuel System Drains … 779

488. FAR 29.1001 Fuel Jettisoning … 78O

**489.-497. RESERVED X11**

### Page 13

7/30/97 SECTION 28. OIL SYSTEM AC 29-2B Page No.

498. FAR29.1011 General … 79l

499. FAR29.1 0130il Tanks … 792

500. FAR29.1 0150il Tank Tests … 792

501. FAR29.10170il Lines and Fitiings … 793

502. FAR29.1019 Oil Strainer or Filter … 793

502A. FAR 29.1019 Oil Strainer or Filter … 794

503. FAR 29.1021 Oil System Drains … 795

504. FAR29.1 023 Oil Radiators … 795

505. FAR29.1 025 Oil Valves … 795

506. FAR29.1 027Transmission and Gearboxes -General … 796

**507.-515. RESERVED SECTION 29. COOLING**

516. FAR 29.1041 General … 8O7

517. FAR 29.1043 Cooling Tests ..................................................................................81 O

517A. FAR29.1043 Cooling Tests … 8l2

518. FAR 29.1045 Cooling Test Procedures … 8l3

519. FAR29.1047 Takeoff Cooling Test Procedures … 8l7

520. FAR29.1049 Hovering Cooling Test Procedures … 8l9

**521-530. RESERVED SEC TION 30. INDUCTION SYSTEM**

531. FAR 29.1091 Air Induction … 83l

532. FAR 29.1093 Induction System Icing Protection … 832

532A. FAR 29.1093 Induction System Icing Protection … 84l

533. FAR 29.1101 Carburetor Air Preheater Design … 842

534. FAR 29.1103 Induction System Ducts and Air Duct Systems … 842

535. FAR 29.1105 Induction System Screens … 845

536. FAR29.1 1071ntercoolers and Afier.Coolers … 845

537. FAR29.1 109Carburetor Air Cooling … 845

**538.-547. RESERVED SECTION 31. EXHAUST SYSTEM**

548. FAR 29.1121 General … 856

549. FAR 29.1123 Exhaust Piping … 857

550. FAR 29.1125 Exhaust Heat Exchangers … 859

**551.-560. RESERVED X111**

### Page 14

SFCTION 32. POWFRP1 ANT CO NTROLS AND ACCESSO RIES Page No.

561. FAR 29.1141 Powerplant Controls: General … 873

562, FAR 29.l142Auxiiiary Power Unit Controls … 874

563. FAR29.1 143 Engine Controls … 875

563A. FAR 29.1143 Engine Controls … 876

563B. FAR 29.1143 Engine Controls … 876

564. FAR29.1145 Ignition Switches … 876

565. FAR 29.1147 Mixture Controls … 877

566. FAR 29.1151 Rotor Brake Controls … 878

567. FAR 29.1157 Carburetor Air Temperature Controls … 879

568. FAR 29.1159 Supercharger Controls … 879

569. FAR 29.1163 Powerplant Accessories … 88O

570, FAR 29.1165 Engine Ignition Systems … 88l

**571.-583. RESERVED SECTION 33. POWERPI A NT FIRE PROTECTION**

584. FAR 29.1181 Designated Fire Zones: Regions Included … 895

585. FAR 29.1183 Lines, Fittings, and Components … 896

586, FAR 29.1185 Flammable Fluids … 897

587. FAR 29.1187 Drainage and Ventilation of Fire Zones … 898

588. FAR 29.1189 Shutoff Means … 899

589. FAR 29.1191 Firewalls … 9OO

590. FAR 29.1193 Cowling and Engine Compartment Covering … 9O4

590A. FAR 29.1193 Cowling and Engine Compartment Covering … 905

591. FAR29.1 1940ther Sutiaces … 9O6

592. FAR 29.1195 Fire Extinguishing Systems … 9O7

593. FAR 29.1197 Fire Extinguishing Agents ...............................................................91 O

594. FAR 29.1199 Extinguishing Agent Containers … 9l3

595. FAR 29.1201 Fire Extinguishing System Materials … 9l4

596. FAR 29.1203 Fire Detector Systems … 9l4

**597.-616. RESERVED SEC TION 34. EQU IPMENT - GENERAL**

617. FAR29.1301 Function and installation … 939

618. FAR29.1303 Flight and Navigation instruments … 939

619. FAR 29.1305 Powerplant instruments … 943

619A. FAR 29.1305 Powerplant instruments … 944

619B. FAR 29.1305 Powerplant instruments … 945

Xlv

### Page 15

7130197 AC 29-2B SECTION 34. EQUIPMENT - GENERAL (continued) Page No.

619C. FAR29.1305 Powerplant instruments … 946

620. FAR29.1307 Miscellaneous Equipment … 946

621. FAR 29.1309 Equipment, Systems, and installations … 947

**622.-631. RESERVED SECTION 35. INSTRUMENTS - INSTALLATION**

632. FAR 29.1321 Arrangement and Visibility … 977

633. FAR 29.1322 Warning, Caution, and Advisory Lights … 978

634. FAR 29.1323 Airspeed Indicating System … 98l

634A. FAR 29.1323 Airspeed Indicating System … 982

635. FAR 29.1325 Static Pressure Systems … 982

636. FAR 29.1327 Magnetic Direction Indicator … 983

637. FAR29.1329 Automatic Pilot Systems … 984

638. FAR29.1331 Instruments Using A Power Supply … 993

639. FAR 29.1333 Instrument Systems … 994

640. FAR 29.1335 Flight Director Systems … 994

641. FAR 29.1337 Powerplant Instruments - … 996

Paragraph (b) - Fuel Gauge Calibration

641A. FAR 29.1337 Powerplant instruments … 997

**642.-651. RESERVED SECTION 36. ELECTRICAL SYSTEMS AND EQUIPMENT**

652. FAR29.1351 Electrical Systems and Equipment .General ................................l OO9

653. FAR 29.1353 Electrical Equipment and installations ...........................................l Ol4

654. FAR 29.1355 Distribution System … 1016

654A. FAR 29.1355 Distribution System ........................................................................l Ol6

655. FAR 29.1357 Circuit Protective Devices … 1017

655A. FAR 29.1357 Circuit Protective Devices ..............................................................l Ol8 656 FAR 29.1359 Electrical System Fire and Smoke Protection ................................l Ol9

657. FAR 29.1363 Electrical Systems Tests … lo2o

**658.-667. RESERVED SECTION 37. LIGHTS**

668. FAR 29.1381 Instrument Lights … 1032

669. FAR 29.1383 Landing Lights … 1032

670. FAR 29.1385 Position Light System Installation … 1033

671. FAR 29.1387 Position Light System Dihedral Angles ..........................................l O33 xv

### Page 16

AC 29-26 SECTION 37. LIGHTS (continued~ 7130197 Page No.

672. FAR 29.1389 Position Light Distribution and intensities ......................................l O33

673. FAR 29.1391 Minimum Intensities in the Horizontal ............................................l O33 Plane of Forward and Rear Position Lights

674. FAR 29.1393 Minimum in Any Vertical Plane of Fomard ....................................l O33 and Rear Position Lights

675. FAR 29.1395 Maximum Intensities In Overlapping Beams .................................l O33 of Forward and Rear Position Lights

676. FAR29.1 397 Color Specifications .......................................................................l O33

677. FAR 29.1399 Riding Light … 1033

678. FAR 29.1401 Anticollision Light System … 1035

**679.-688. RESERVED SEC TION 38. SAFETY EQUIPMENT**

689. FAR 29.1411 General … 1044

690. FAR 29.1413 Safety Belts: Passenger Warning Device .....................................l O46

691. FAR29.’l4l5 Ditching Equipment .......................................................................l O47

**692. RESERVED**

693. FAR 29.1419 Ice Protection .................................................................................l O5O

**694.-701. RESERVED SECTION 39. MISCELLANEOUS EQUIPMENT**

702. FAR29.1431 Electronic Equipment .....................................................................l O76

703. FAR29.1433 Vacuum Systems ...........................................................................l O76

704. FAR29.1435 Hydraulic Systems .........................................................................l O77

705. FAR29.1439 Protective Breathing Equipment ....................................................l O8O

706. FAR29.1457 Cockpit Voice Recorder .................................................................l O8l

707. FAR29.1459 Flight Recorders ............................................................................l O83

708. FAR 29.1461 Equipment Containing High Energy Rotors ...................................l O85

**709.-717. RESERVED SECTION 40. 0 PERATING LIMITATIONS**

718. FAR29.1501 General ..........................................................................................l O96

719. FAR29.1503 Airspeed Limitations: General ......................................................l O96

720. FAR29.1505 Never-Exceed Speed ....................................................................l O98 721, FAR29.1509 Rotor Speed ..................................................................................ll Ol

722. FAR 29.1517 Limiting Height-Speed Envelope ...................................................ll O3

723. FAR29.1 519Weight and Center of Gravity ........................................................ll O4

724. FAR 29.1521 Powerplant Limitations ..................................................................ll O4 xv\

### Page 17

SECTION 40. 0 PERATING LIMITATIONS (Continued) Page No.

724A. FAR 29.1521 Powerplant Limitations ..................................................................ll O5 724B. FAR 29.1521 Powerplant Limitations ..................................................................ll O6

725. FAR29.1 522 Auxilia~Power Unit Limitations ....................................................ll O7

726. FAR 29.1523 Minimum Flightcrew .......................................................................ll O7

727. FAR29.1525 Kinds of Operations … llll

728. FAR 29.1527 Maximum Operating Altitude … lll2

729. FAR29.1 5291nstructions for Continued Aiwodhiness … lll2

(Maintenance Manual)

729A. FAR 29.1529 Instructions for Continued Aimofihiness … lll5

(Maintenance Manual)

**730.-739. RESERVED SECTION 41. MARKINGS AND PLACARDS**

740. FAR29.1541 General (See Paragraph 781) … ll26

741. FAR 29.1543 Instrument Markings: General (See Paragraph 781) ....................1 126

742. FAR 29.1545 Airspeed Indicator (See Paragraph 781) … 1126

743. FAR 29.1547 Magnetic Direction indicator … ll26

744. FAR 29.1549 Powerplant Instruments (See Paragraph 781) … 1126

745. FAR 29.1551 Oil Quantity indicator … ll26

746. FAR 29.1553 Fuel Quantity indicator … ll27

747. FAR 29.1555 Control Markings … ll27

748. FAR 29.1557 Miscellaneous Markings and Placards … ll29

749. FAR 29.1559 Limitations Placard … ll3O

750. FAR 29.1561 Safety Equipment ...m … ll3l

751. FAR29.1 565 Tail Rotor … ll32

**752.-761. RESERVED SECTION 42. ROTORC RAFT FLIGHT MANUAL**

762. FAR 29.1581 General … ll43

763. FAR 29.1583 Operating Limitations … ll5l

763A. FAR 29.1583 Operating Limitations … ll54

764. FAR 29.1585 Operating Procedures … ll54

765. FAR 29.1587 Performance Information … ll56

765A. FAR 29.1587 Performance information … ll6O

766. FAR 29.1589 Loading information … ll6O

767.-774. RESERVED Xvll

### Page 18

**CHAPTER 3- MISCELLANEOUS AIRWORTHINESS STANDARDS**

775.

776.

777.

778.

779.

780.

781.

781A.

782.

783.

784.

785.

786.

787.

788.

Page No.

IFR Certification … ll7l

Certification Procedure for Rotorcraft Avionic Equipment … ll92

Standardized Test Procedure for Rotorcraft DC … 1209

Electrical System Tests

Standardized Test Procedure for Rotorcraft Generator Cooling … 1214

FARs29.1301, 29.1309, and29.1322 Annunciator Panels … l2l7

**RESERVED**

FARs29.1541, 29.1543, and 29.1549 instrument Markings … l22O

FAR 29.1549 Powerplant Markings … l224

Rotorcraft and Systems Certification for Category II … 1224

Operations

Structural Condition indicators … l23O

Full Authority Digital Electronic Controls (FADEC) for … 1234

Installations with Category A Engine Isolation

Agricultural Dispensing Equipment installation … l236

Emergency Medical Service (EMS) Systems, Installations , … 1242

Interior Arrangements, and Equipment

**RESERVED**

Substantiation of Composite Rotorcrafi Structure … l262

Appendix 1 Fatigue Evaluation of Transport Category Rotorcraft … AI-I

Structure (Including Flaw Tolerance)

Appendix 2 Rotorcraft OEI Power Assurance … A2-I

Appendix 3 Advisory Material for Compliance with Rotorburst Rule … A3-l

Xvlll

### Page 19

ILLUSTRATIONS The first digits in the figure or table number are the same as the related paragraph number.

The digits after the hyphen are the number within the paragraph. Most figures or tables are located at the end of the related paragraph.

Fiaure/Table No. m 55-1 H-V Requirements 55-2 IGE Controllability Requirements 55-3 Shaft Horsepower vs. Turbine Outlet Temperature - Sea Level Standard Day 55-4 Uninstalled Takeoff Power Available 55-5 Installed Takeoff Power Available, Anti-ice, 400 RPM 55-6 Power Assurance Check Chart 58-1 Takeoff Performance Category A 58A-1 Takeoff Performance Category A 60-1 Category A Vertical Takeoff Profile Ground Heliport 60-2 Category A Vertical Profile Pinnacle 60A-I Category A Vertical Takeoff Profile Ground Level Heliport 60A-2 Category A Vertical Takeoff Profile Elevated Heliport 64-1 Conventional Takeoff Profile Category B 68-1 Autorotational Characteristics - Typical 70-1 Category A Conventional Landing - Clear Heliport 75 76 77 78 79 80 87 89 99 100 110 111 117 128 139 X1X

### Page 20

AC 29-2B Fiaure~able No. a 70-2 70A-1 70A-2 72-1 72-2 85-1 110-1 137-1 140-1 140-2 162-1 162-2 230-1 249-1 249-1 (Table 1) 328-1 337-1 373-1 386-1 xx Category A Vertical Landing Category A Conventional Landing - Clear Heliport Category A Vertical Landing Height-Velocity Diagram Altitude/Weight Accountability Static Longitudinal Stability Demonstration Points Load Factor Gross Weight Curve Sample Yaw/Forward Speed Diagram Sample Yaw/Forward Speed Diagram (View Looking Forward) (View Looking Forward) Comparison Percent of Time in Level Flight Example of Casting Load Sheet Retract Actuator Support-Landing Gear Castings to be Static Tested for Model XYZ Cockpit Visibility Sea State Code - World Meteorological Organization Summary of Part 133 Rotorcraft Load Combinations Certifiable Under $29.865 Continuous Icing-Temperature vs. Altitude Limits 7/30/97 u 140 150 151 160 161 183 214 241 246 248 275 276 348 381 383 476 501 558 592

### Page 21

7/30/97 Fiaure/Table No. m 386-2 386-3 386-4 386-5 423A-1 423A-2 423A-3 425-1 425-2 447-1 447-2 447-3 447-4 447-5 447-6 Intermittent Icing-Temperature vs. Altitude Limits Mean Effective Drop Diameter-Microns Intermittent Icing-Liquid Water Content vs.

Drop Diameter Supercooled Clouds Rotor Drive System Endurance Takeoff Run Portion of Rotor Drive System Endurance Test If 2 % Minute OEI Rating is Requested.

Takeoff Run Portion of Rotor Drive System Endurance Test If 30 Second/2 Minute OEI Rating is Requested.

Cantilevered Shaft, First Critical Speed Shaft Between Support Bearings, First Critical Speed Static Tension and Shear Loading Modes Hose Assembly Tests Minimum Average and Individual Loads for Hose and Hose-End Fitting Combinations Typical Method of Breakaway Fuel Fitting Load Calculations Typical Methods of Frangible or Deformable Attachment Load Calculations Typical Methods of Frangible or Deformable Attachment Load Calculations AC 29-2B w 593 594 595 596 660 661 662 675 675 719 720 721 722 723 724 Xxl

### Page 22

AC 29-2B Fiaure/Table No. m 450A-I 516-1 518-1 589-1 637-1 637-2 693-1 693-2 693-3 693-4 775-1 782-1 782-2 785-1 785-2 786-1 Xxll Fuel Flow Additional Cooling Test Point All-Engine-Critical Altitude Additional Cooling Test Point Table of Materials and Gages Acceptable for Fireproof Protective Devices with Flat Surface Geometries , Deviation Profile Operational Limitation Continuous Icing-Temperature vs. Altitude Limits Intermittent Icing-Temperature vs. Altitude Limits Continuous Icing-Liquid Water Content vs Drop Diameter Intermittent Icing-Liquid Water Content vs.

Drop Diameter Rotorcraft Dynamic Stability Requirements for IFR Category II Approach Evaluation Category II Approach Evaluation (cont.)

Acceptable Ultimate Load Factor for Agricultural Dispensing Equipment Design Sketch of Tank Pressure Test Typical Liquid Oxygen System 7130197 a 734 809 816 903 990 992 1064 1065 1066 1067 1191 1228 1229 1237 1241 1259

### Page 23

Figure/Table No. m 786-2 Typical Liquid Oxygen System - Using Combination Valve APXI-I S-N or P-N Curve Usage APXI-2 Crack Growth for Single Element Structure APXI-3 Crack Growth for Remaining Elements of Multiple Element Structure APX2-I Installed Engine Fuel Flow Capacity Check APX3-I Estimated Path of Fragments APX3-2 Single One-Third Disc Fragment APX3-3 Intermediate and Small Pieces of Debris APX3-4 Blade Fragment Definition APX3-5 Distribution of Translational and Rotational Kinetic Energy of Rotor-Component Fragments as a Function of Fragment Size 8 APX3-6 Cross Section through Aircraft at Plane of Rotation of the Engine Disk Fragment AI-15 AI-18 AI-20 A2-9 A3-I O A3-I 1 A3-12 A3-I 3 A3-14 A3-I 5 Xxlll

### Page 24

AC 29-26 Abbreviations and Definitions AC advisory circular ACO aircraft certification office ADF automatic direction finding ADI attitude direction indicator AEO all engines operating AFCS automatic flight control systems AGL above ground level AHRS attitude heading reference system Amdt. Amendment APU auxiliary power unit ATC air traffic control BIM blade inspection method CAM cockpit area microphone CAR Civil Air Regulations CAS calibrated air speed CBIM cockpit blade inspection method CDP critical decision point CG center of gravity CPS cycles per second CRFS crash resistant fuel system CRT cathode ray tube CVR cockpit voice recorder DER designated engineering representative DME distance measuring equipment ECAS engine caution advisory systems ECU environmental control unit EFIS electronic flight instrument system EFP engine failure point EHE exhaust heat exchanger ELT emergency locator transmitter EMI electromagnetic interference EMS emergency medical service EOL end of life FAA Federal Aviation Administration FADEC full authority digital engine control FAR Federal Aviation Regulations FMEA failure mode and effects analysis 7/30/97 FPM feet per minute FPS feet per second GPS global positioning system HF high frequency HIRF high intensity radiated fields H*RD high rate of discharge HUMS health and usage monitoring systems HV height-velocity IAS indicated airspeed ICAO International Civil Aviation Organization ICS inter-communication system IFR instrument flight rules IGE in ground effect IIDS integrated instrument display system ILS instrument landing system IMC instrument meteorological conditions INS inertial navigation system ISA international standard atmosphere ISIS integral spar inspection system ITT inter-turbine temperature KCAS knots calibrated airspeed KIAS knots indicated air speed KTAS knots true airspeed LDP landing decision point LWC liquid water content MCP maximum continuous power MGT measured gas temperature MSL mean sea level MVD median volume diameter NDI non-destructive inspection NM nautical mile NPRM notice of proposed rulemaking NTSB National Transportation Safety Board NVG night vision goggles OAT outside air temperature OEI one engine inoperative OGE out of ground effect PBA pitch bias actuator Xxiv

### Page 25

PCF post crash fire PIO pilot induced oscillation PSIG pounds per square inch gauge QPL qualified parts list RFM rotorcraft flight manual RFMS rotorcraft flight manual supplement RPM revolutions per minute RTCA Radio Technical Commission of Aeronautics RVR runway visual range SAE Society of Automotive Engineers SAS stability augmentation system SCAS stability and control augmentation systems SRM structural repair manual STC supplemental type certificate STOL short takeoff and landing TBO time between overhaul TC type certificate TCDS type certificate data sheet TDP takeoff decision point TIA type inspection authorization TIR type inspection report TOT turbine outlet temperature TSO: technical standard order TVP true vapor pressure VBIM visual blade inspection method VFR visual flight rules VMC visual meteorological conditions VOR very high frequency omnidirectional range radio VSI vertical speed indicator V/STOL vertical/short takeoff and landing VTOL vertical takeoff and landing WAT weight, altitude, temperature Altitudes HD density altitude HP pressure altitude V speeds VD diving speed VH speed in level flight with maximum continuous power VMO maximum operating limit speed VNE never-exceed speed VTOSS takeoff safety speed for Category A rotorcraft VX speed for best angle of climb VY speed for best rate of climb MMO maximum operating math number N speeds NF free turbine speed NG gas generator speed NP power turbine speed NR rotor speed Coefficients CD coefficient of drag CL coefficient of lift CP coefficient of power CT coefficient of thrust xxv

### Page 26

TF P~ mendm,ent 21 5(IJ.

1.-2. ~.

Intentionally Left Blank Par 1 I (thru 3)

### Page 27

AC 29-2B

**3. RESERVED.**

4 7/30/97 intentionally Left Blank Par 3

### Page 28

4. ~ 21.16 SPECIAL CONDITIONS.

a. The Process. Chapter 2, Section 1, Paragraph 8 of the Type Certificate Handbook, Order 8110.4, 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-I 10. That otfice 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-I 00 will issue the special conditions.

b. Basis for Develo~ment.

(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.

Par 4 5

### Page 29

AC 29-2B

**5.-7. RESERVED.**

6(thru 8) 7/30/97 Intentionally Left Blank Par 5

### Page 30

8. 671.3 1 PF DFS lG~. 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.

Par 8 9

### Page 31

AC 29-2B

9.-11. ~.

10(thru 12) 7/30/97 Intentionally Left Blank Par 9

### Page 32

12. ~ 21.33 INSPFCTIO N AND TESTS.

a. Armlicant Res~ onsibilitv. 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 FANAUTHORITY for testing.

(2) Conduct all inspections and tests necessary to determine compliance with the airworthiness and noise requirements.

b. FWVAUTHOR ITY Respo nsibilitw

(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.

### Page 33

AC 29-2B

**13.-15. RESERVED.**

14(thru 16) 7/30/97 Intentionally Left Blank Par 13

### Page 34

7130197 AC 29-2B

16. $21.35 (Amendment 21-59) FLIGHT TESTS.

a. Exdanation.

(1) This section outlines the requirements of the applicant for aircraft type certification and should be used in conjunction with Order 8110.4, 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 FAWAUTHORITY 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 FANAUTHORITY verification flight test should include, but not be limited to, the critical or marginal results contained in the applicant’s flight test report.

The FANAUTHORITY’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 FANAUTHORITY can provide technical flight test 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 FAIVAUTHORITY engineering flight tests are necessary.

These tests are normally specified in the Type Inspection Authorization (TIA). At the same time the FANAUTHORITY 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 24).

(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 Groupinas.

### Page 35

(i) Weight and e.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) Spe cial 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.

b. Procedures.

(1)Type Certification Flight Tests.

(i) Prior to initiating official FAWAUTHORITY 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 FANAUTHORITY, 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) Certification Handbook 8110.4, Paragraph 67, contains a requirement that the applicant must keep the FANAUTHORITY 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 maybe 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 FANAUTHORITY manufacturing inspector and certification engineer are aware of all configuration changes found necessary as a result of FANAUTHORITY 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.

18 Par 16

### Page 36

(v) Additional flight test responsibilities, procedures, and requirements during the certification flight test process are contained in Certification Handbook 8110.4, 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) Function and reliability (F&R) testing must be accomplished on an aircraft which is in conformity with the approved production configuration. F&R testing should follow the type certification testing described in Paragraph 16b(l ) above to assure that significant changes resulting from type certification tests can be 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 emphasize systems, operating conditions, or environments found particularly marginal during type certification tests.

(vi) See Handbook 8110.4, Paragraph 166(c), for additional information and procedures.

Par 16 19

### Page 37

AC 29-2B

**17.-23. RESERVED.**

ZO(thru 26) 7/30/97 Intentionally Left Blank Par 17

### Page 38

24. S 21.39 (Amendment 21-5 9) FLIGHT TEST INSTRUMENT CALIBRATION AND CORRECTION REPORT.

a. Exdanation. 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) established Standards.

(3) Regardless of the record form, the accuracy of the record must be by reference to a laboratory standard traceable to the National Bureau of 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. Procedu res. Prior to conducting flight tests, the FANAUTHORITY 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 gaged components. Also, environmental effects such as vibration, humidity, temperature, etc., should be considered when determining whether recalibration is necessary.

(2) The highest and lowest magnitude of the parameter being recorded should be considered when establishing the scale for instrumentation. Ideally, the highest magnitude throughout the flight would fall on the maximum indicating point of the recording.

Par 24 27

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4C 29-2B

**25.-30. RESERVED.**

28 (thru 34) 7/30/97 Intentionally Left Blank par 25

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**CHAPTER 2. PART 29**

**AIRWORTHINESS STANDARDS**

**TRANSPORT CATEGORY ROTORCRAFT SECTION 1. GENERAL**

31. 529.1 (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 multiengined 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 maybe single or multiengined 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 multiengined 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 Enaine Isolation. Categoty B multiengined 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 Cert ification, Categories A and B. A multiengined 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.

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b. Procedures. None.

31A.~~ APPLIC BILITY.

a. Explanation. Amendment 29-39 revised the reference in fj 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 multiengined 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 multiengined 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 Enaine Isolation. For single engine rotorcraft and multiengined 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) ~ Isolation. Category B multiengined 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) ~ ifi ion ri nd B. A multiengined 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 31 does not apply to rotorcraft certified with Amendment 29-39 or later.

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7/30197 AC 29-2B

32. ~ 29.2 (Amendment 29 -32) SPECIAL RETROACTIVE REQU IREMENTS.

a. Exdanation.

(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 Cl 14-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 Cl 14 or approved under equivalent standards such as those contained in Part 29.

(4) Load Distribution and Des ian 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 335b(l )(i) (~ 29.785) of this AC.

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(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.

(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 aircrafi specification or TCDS.

(2) Structural analysis or static test maybe 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 Cl 14 approval does not constitute approval for installation of a restraint system in a rotorcraft design subject to dynamic tests.

(i) AC 20-137 dated March 30, 1992, 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.Og, and for later designs (Amendment 29-29 and later), it is 8.Og.

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7/30/97

**33.-41. RESERVED.**

Intentionally Left Blank AC 29-2B 3g(thru 47)

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**AC 29-2B SECTION 2. FLIGHT - GENERAL 7130197**

42. ~ 29.21 (Amendment 29-24) PROOF OF CO MPLIANCE.

a. Exp Ianation.

(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 FWAUTHORITY-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 FAWAUTHORITY 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 FAWAUTHORITY 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. FAWAUTHORITY compliance maybe 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 FAWAUTHORITY flight tests. The FANAUTHORITY 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 petiormance 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 (lGE) and, if required, out-of-ground effect (OGE) conditions. From these data the hover ceiling may be calculated.

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(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 5$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. Procedu res.

(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 reviewed, a Type Inspection Authorization, or equivalent, can be issued. The FANAUTHORITY 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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42A. ~~ F OF C MPLIANCE.

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 42 continues to apply.

43. 62 9.25 (Amendment 29-12) WEIGHT LIMITS.

(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) Jett isonable External CaraQ.

(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 50 Par 42A

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approval whenever the “standard” structural limitations are exceeded (Review Case Nos. 37 and 55).

(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 FAWAUTHORITY 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 testinstrumentation 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. However, 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 762 of this AC concerns the RFM and its contents.

(3) See Paragraph 230, ~ 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.

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44. ~ . n~ RAVITY IMITS.

a. ~xrianation.

(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 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 52 Par 44

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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.

45. 629,29 (Amendment 29 -15) EM PTY WEIGHT AND CORRESPONDING CENTER OF GRAVITY.

a. ~xdanatio~. The empty weight of the rotorcraft consists of the airframe, engines, and all items of operating equipment that have fixed locations and are permanently installed in the rotorcraft. It includes fixed ballast, unusable fuel, and full operating fluids except water intended for injection in the engines.

(1) Fixed ballast refers to ballast that is made a permanent part of the rotorcraft as a means of controlling the certificated empty weight CG.

(2) Compliance with Paragraph (b) of S 29.29 is accomplished by the use of an equipment list which defines 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 a manufacturing inspector and is normally made on a production rotorcraft rather than a prototype. If a manufacturer wishes to avoid weighing each production rotorcraft and has been issued a production certificate, the manufacturer may make a detailed proposal defining the procedures used 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 *I percent on empty weight and *0.2 inches on CG After this procedure is established, the empty weight and CG maybe 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 AC 91-23A, Pilots Par 44 53

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AC 29-2B 7130/97 Weight and Balance Handbook, June 9, 1977, for a sample weight and balance procedure.

(3) The weight and balance should be recalculated if a modification (or series of modifications) to the rotorcraft results in a significant change to the empty weight.

Additionally, this change in empty weight should be reflected with the weight and balance information contained in the Rotorcraft Flight Manual (RFM) or Rotorcraft Flight Manual Supplement (RFMS).

c. Ballast Load ina and Tvpe.

(1) Ballast loading of the rotorcraft can be accomplished in any manner to achieve a specific CG location. It is acceptable for such ballast to be mounted outside the physical confines of the rotorcraft if the flight test objectives are not affected by this arrangement. In flight test work, loading problems will occasionally be encountered in which it will be difficult to obtain the desired CG limits. Such cases may require loading in engine compartments or other places not designed for load carrying. When this condition is necessary, care should be taken to ensure that local structural stresses are not exceeded or that the rotorcraft flight characteristics are not changed due to increased moments of inertia by attaching the ballast to extreme CG locations which may not be designed for the added weight.

(2) Two types of ballast that maybe used in loading are solids or liquid. The solids are usually high density materials such as lead while the liquid usually used is water. In critical tests, the ballast may be loaded in a manner so that disposal in flight can be accomplished. In any case, the load should be securely attached in its loaded position so shifting or interference with safety of flight will not result.

46. ~ 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 385 of this advisory circular for information on ballast provisions.

b. Procedures. None.

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47. 5 29.33( Amendme nt 29-15) MAIN ROTOR SPEED AND PITCH LIMITS.

a.

(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 suficient to maintain autorotational rotor speed values within limits. If this occurs, the manufacturer may elect to adjust the low pitch stops as a maintenance procedure at extreme ambient conditions provided the flight and maintenance manuals clearly present the rigging requirements and procedures. There must be sufficient “overlap” of ambient conditions between configurations such that rerigging is not required whenever ambient temperature and surface elevation change slightly. Any 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.

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AC 29-2B 7’130197

(4) Additional RPM Ranaes. 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 FIW/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 multiengined rotorcraft where there is not an automatic increase in remaining engine(s) power output upon failure of an engine. Although today’s multiengined 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 beat 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 721 (~ 29.1509) for additional information on rotor limits determination and testing.

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7/30/97

**48.-54. RESERVED.**

intentionally Left Blank AC 29-2B 57 (thru 63)

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AC 29-2B sECTION 3. PERFORMANCE 7/30/97 55, ~~ -24 RAL.

a. Exp Ianation.

(1) Amendment 29-39 adopts new and revised airworthiness standards for the performance of transport category rotorcraft. As part of this change, several sections within the performance section of Subpart B were renumbered. The performance section of this guidance material has been organized for easy use with rotorcraft certified 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 pettinent information can be found.

(2) Section 29.45 lists some of the rules and standards under which the performance requirements are to be met. This paragraph will provide general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables which affect the flight test results that must be taken into account. Some of these variables are wind, temperature, altitude, humidity, rotorcraft weight, power, rotor RPM, center of gravity, etc. The test results should be analyzed and expanded by approved methodology within the guidelines of this paragraph. 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.

(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 height-velocity (HV) determination, low rotor speeds have been authorized based upon additional structural and drive train substantiation (see Paragraph 47).

(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.

b. Procedures

(1) Winds For Te stinq.

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(i) Allowable wind conditions will vary with the type of test and will also bedifferent fordifferent types andgross weight rotorcrafi. Forexample, higher winds can usually be tolerated for takeoff and landing distance tests than for hover performance. Likewise higher winds can sometimes be tolerated during hover performance testing on large, heavy rotorcraft with high rotor downwash velocities than for smaller rotorcraft with rotor downwash velocities. Generally, unless the effects of wind on hover performance tests can be determined and/or accounted for, hover performance testing should be conducted in winds of 3 knots or less.

(ii) Past experience has shown that a @eadv wind of Oto 10 knots will result in acceptable takeoff and landing petiormance 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, the same general wind criteria used to obtain hover performance; i.e., up to 3 knots, should be adhered to for vertical performance determination. In actuality, a rotorcraft may exhibit reduced IGE hover performance in winds 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., Oto 3 knots.

(iii) As can be seen from the foregoing, there is no such thing as an exact allowable wind for a particular test or rotorcraft. The flight test team must decide on the allowable wind for each condition based on all available information and their engineering judgment. The following summary of allowable wind conditions are 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 3 knots (D) Height-velocity -0 to 3 knots

(iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests. 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. Using F/lA/AUTHORITY-approved methodology, hover, takeoff, and landing, performance may be extrapolated and/or interpolated from test data up to a maximum of *4,000 feet. Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics should not be Par 55 65

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extrapolated more than approximately 2,000 feet density altitude from the test altitude.

Cruise stability/controllability tests should be evaluated at least at two different altitudes, the lowest practical altitude and approximately the highest cruise altitude requested for approval. This can allow an interpolation of approximately 10,000 feet. As in all testing, extrapolation and/or interpolation should only be considered if all available information and engineering judgment indicate that regulatory compliance can be met at the untested conditions.

(3) A~.

(i) Exp Ianation.

(A) Two altitudes are normally presented in the RFM to define the operating envelope of a rotorcraft:

- Maximum operating altitude; and,

- Maximum takeoff and landing altitude.

(B) Maximum operating altitude, is an operating limitation required by $29.1527 and delineates the maximum altitude up to which operation is allowed.

This altitude normally constitutes the maximum cruise or enroute altitude.

(C) Maximum weight, altitude and temperature for takeoff and landing constitutes a limitation. The maximum takeoff and landing altitude may be coincident with but never above the maximum operating altitude limitation. Takeoff and landing and hover ceiling data and presentation requirements are presented in ~~ 29.51,29.53,29.59,29.63, 29.73,29.1583 and 29.1587.

(ii) Procedures.

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

(1) Takeoff ($S 29.51-29.63)

(2) Climb (~~ 29.64-29.67) (Q) Performance at minimum operating speed (S 29.49)

(q) Landing ($ 29.75)

(5) Limiting height-speed envelope (~ 29.87) 66 (Q) IGE controllability ($ 29.143c) Par 55

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7130197 AC 29-2B (Z) Cooling ($S 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 AC.

(B) As detailed in subparagraph b(2) above, the maximum allowable extrapolation of H-V, IGE controllability and engine operating characteristics is *2,000 feet. Therefore, the maximum takeoff and landing altitude presented in the RFM is not normally more than 2,000 feet above the density altitude experienced at the high altitude test site.

(C) Prior to Amendment 29-21, H-V information was an operating limitation. With the adoption of Amendment 29-21, the H-V 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 55-1 and 55-2.

(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 subparagraph A of this paragraph.

(i) 13ackground.

(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.

Par 55 67

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(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 math 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/or materials. Therefore, to satisfy S 29.1309(a), the applicant needs to substantiate the total rotorcraft at the extreme temperatures for which certification is requested.

(ii) Procedures.

(A) The FAPJAUTHORITY 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, V~~/V~ evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.

(B) Extrapolation of test data should only be allowed if the applicant’s predicted or calculated data is verified by actual test but in any case extreme caution should be used for extrapolations that are -1 O°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 only along an established W/a line within the allowable altitude extrapolation range.

(6) Engine Power - Turboshaft Enaine.

(i) Background.

Par 55

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7130/97 AC 29-2B (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/or speeds. The specification values may be either a -or w. Some engine manufacturers certify an engine to a specified horsepower at a particular engine temperature or speed rating with higher allowable limits. The W 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 55-3 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 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 55-7.

(E) Several power losses may be associated with installing an engine in a rotorcraft. Typical losses are air inlet losses, gear losses, air exhaust losses, and powered accessory losses such as electrical generators. Additional flight manual performance considerations are the torque indicating system accuracy and torque 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 Par 55 69

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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 and/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 55-8 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 and/or RPM can significantly decrease for a period of time as torque is held constant.

Said another way, torque will increase if RPM and/or temperature is heId constant. This is a characteristic typical of turbine engines due largely to expansion of turbine blades and reduced clearances in the engine. Some engines may show a temperature increase at constant power due to engine or temperature sensing system peculiarities.

An engine will usually establish a stabilized relationship of power parameters in approximately 2 or 3 minutes. For this reason, the following procedure should be used when obtaining in-flight engine data.

(1) To determine the applicable value (takeoff, 30-second, and 2 l/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 l/2-minute power levels as necessary. Record the engine parameters as soon as the specification 70 Par 55

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torque, temperature, orspeed is attained. Care must betaken nottoexceed a limit.

These readings should be obtained approximately 15 seconds aller power is initially applied.

(z) To determine the 30-minute and/or 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 l/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 ratinq with higher maximum m value approved. Only the rating values should be used to determine the installation losses.

The w values of engine temperature and/or 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 multiengined rotorcraft, the engine location may result in different installation losses between engines. If this condition exists, multiengined performance should be based on a total of the different minimum installed specification horsepower values. One engine inoperative performance must be based on the loss of the engine which has the lowest installation losses. Additionally, the power losses due to such items as accessory bleed air, particle separators, etc., must be accounted for accordingly.

(E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical NG speed to a constant corrected speed at cold temperatures. Others may limit power to a 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.

Par 55 71

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(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 55-9 is a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the rotorcraft flight manual performance. Some power assurance curves have maximum allowable N~ 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) Deteriorate d Engine Power - Turboshaft En sine.

(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 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:

72 Par 55

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(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/aircraft installation and fuel scheduling mechanism.

(z) Performance of customer bleed air systems maybe degraded slightly. No problem would be anticipated unless certain items within the system depend on a critical Pc for their function.

(a) The maximum attainable gas producers speed, and thus power available under certain ambients, may be affected if Pc pressure is an input to the fuel scheduling mechanism.

(4) Systems for surge protection which schedule on Pc 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 Pc because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings.

(ii) Procedures, (A) The need for flight tests to verify predicted power available with deteriorated engines depends on the scope of testing which occurred during initial certification. If the original rotorcraft certification included flight testing as described in Paragraph (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 (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 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.

Par 55 73

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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 (6) (engine power-turboshaft engines) herein. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure already in place for the specification engine, which could eliminate the need for flight test evaluation.

(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 and/or the operator has demonstrated his ability to operate safely with deteriorated engines.

55A. 629.45 (Amendment 29-24) GENERAL.

a. ~planat ion. 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 rotorcraft flight manual (RFM).

b. Procedures. All of the policy material pertaining to this section remains in effect. In addition, the power assurance information included in the RFM should be verified. Although this requirement is normally met with a power assurance curve, other methods of compliance may be proposed.

55B. ~ 29.45 (Amendment 29-24) GENERAL.

a. Fxp anatI ion. Amendment 29-34 added the requirements for certification of 30-second/2-minute One Engine Inoperative (OEI) power ratings. For rotorcraft approved for the use of 30-second/2-minute OEI, partial power checks currently accomplished with approved power assurance procedures for lower power levels may not be sufficient to guarantee the ability to achieve the 30-second power level.

b. Procedures. Information provided in Appendix 2 of this AC includes guidance material on power assurance procedures to ensure that the OEI power level can be achieved. The guidance material presented in Paragraphs 55 and 55A continue to apply.

74 Par 55

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AC 29-2B 7130/97 76 Par 55

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600 500 400 300 200 100 0 ENG S/N 1234 / MIN CERTIFIEDSPECIFICATION ENGINE 1000 1100 1200 1300 1400 1500 CORRECTED TOT “F

**SHAFT HORSEPOWER VS TURBINE OUTLET TEMPERATURE -**

SEA LEVEL STANDARD DAY Par 55 77

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2( It 16 14 12 10 8 6 4 2 0 200 300 400 500

SHAFT HORSEPOWER AVAILABLE FIGURE 55-4. UNiNSTALLED TAKEOFF POWER AVAllJIBLE 78 Par 55

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so0 2( 18 1( 1, 12 1- W LLl IL 1 lC 8 f 4 2 0 — I , I I I I 1 —1——, I –1 I I I

- –1 1 I I— I i —1 I 1 !

J I ~ I I I I I I I , I I I I I I I I I I 1 I I I I I I I I —r , I L I \ I I ~— I I – l– I I I I _i _. “–1— ‘- :– — –“i-– I 1!

r —1—— I I I I I I I z’ — I I I I — I I ( I I I— —~ — —[ I I I I I — I I I I I I I 1’ I I I I I 200 300 400

SHAFT HORSEPOWER AVAILABLE ..l— —+ I I I !

I— –L 1– I I —l— I I I 1 –f — I -l _ I –+––l– I I I I FIGURE 55-5. INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM Par 55 79

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2UI N n<Cj o 0 I 80 Par 55

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56. ~ 29.49 (Amendment 29-39) PERFORMANCE AT MINIMUM OPERATING SPEED. HOVER PERFORMANCE FOR ROTORCRAFT.

(For petiormance at minimum operating speed and for hover performance prior to Amendment 39, see S 29.73 and Paragraph 69.)

a. Explanation.

(1) Amendment 29-39 redesignated ~ 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 IGE for Categoty B rotorcraft. Additionally, 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 1%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 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 Paragraph 57 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 81

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to measure the rotorcraft’s pull on the cable. Hover heights are based on skid or wheel height above the ground. During tethered hover tests, the rotorcraft should beat 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 maybe 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 torques.

(ii) Rotor speed.

(iii) Ambient temperatures.

(iv) Pressure altitude.

(v) Fuel used (or remaining).

(vi) Load cell reading.

(vii) Generator(s) load.

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As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in Paragraph 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 F/AA/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.

57. ~ 29.51 TAKEOFF DATA - GENERAL.

a. Exdanation. Section 29.51 details the conditions under which takeoff performance data can be obtained and presented in the FANAUTHORITY 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 Par 56 83

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limitations. The recommended technique used to take off under such conditions is to accelerate the rotorcraft in-ground-effect (lGE) 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. Procedu re. 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 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).

57A. ~2 1. m m~L.

Explanation. Amendment 39 added takeoff requirements in new ~~ 29.55,

29.60a”29.61 and 29.62.

84 Par 57

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7/30197 AC 29-2B

b. Procedures. The guidance material presented in Paragraph 57 continues to apply.

58, fj 29.53 TAKEOFF: CATEGO RY A.

a. Exdanation.

(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 FANAUTHORITY 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 (V~o~~), 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 Par 57A 85

### Page 77

AC 29-2B 7/30197 pedals). Manipulation of thethrottle 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.

(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 sutface. 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 58-1.

b. Procedures.

None.

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7130/97 AC 29-2B I 00 HEIGHT ABOVE GROUND (FEET) 34 /’I. ..-— - ——-. —--- — —.. —-—- ——. — -—- &’ .$+ --.—- - --—- —-- —t-- 1 I I

ALL ENGINES OPERATING 1= El lRANSITION SEGMENT + *TAKEOFF CLIMB SEGMENT+ I

TAKEOFF SEGMENT RTO SEGMENT ‘~ ~ DEMONSTRATEDSAFEOEI LANDINGS -—~

**HORIZONTAL DISTANCE FIGURE 58-1. TAKEOFF PERFORMANCE CATEGORY A I 87**

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58A. ~~ n AKEOFF: CAT RY A,

a. Exp Ianation. 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 58A-1.

b. Procedures. None.

par 58A

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7/30/97 100oft.

I~ 9 200ft.

3 t H-V AVOID (29.87) vv ~ l~fprn< =’~E “F EFPr 35ft.

Is ft. ‘“ RejectedTO. Dist (29,62) --+ 1 [ + T.O.Dist (2?3.61) +’ Horizontal Distance AC 29-2B vv~’lmfpm (Z9.67a.2) Iocoft..— LEVEL ACCEL TOVY FIGURE 58A-1. TAKEOFF PERFORMANCE CATEGORY A Par !j8/+ 89

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59. ~ 29.55 (Amendme nt 29-39). TAKEOFF DECISI ON POINT: CATEGORY A.

a. Exdanation.

(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.

60. ~ 29.59 (Amendment 29 -24) TAKEOFF PATH: CATEGORY A.

a. Exdanation. 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 58-1, Paragraph 58 of this advisory circular.) CDP is a “go/no-go” condition which is analogous to VI speed in transport airplanes. Prior to 90 Par 59

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7/30197 AC 29-2B CDP the pilot is “stop” oriented, 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) JRejected 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) Takeo ff 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 VToss 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 %-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 Par 60 91

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AC 29-2B 7130197 above the surface, and complete landing gear retraction prior to exhausting the 2 Y2-minute time limit.

(ii) The takeoff safety speed, V~o~~, is a speed at which 100 FPM rate of climb is assured under conditions defined in S 29.67(a)(l). The takeoff distance is the distance from initial hover to the point at which V~o~s 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 V~o~s 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~oss to VY in a continuous maneuver without losing altitude, including any nonfigurative 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 sutface area required for takeoff. If takeoff conditions scheduled include considerable surface area (on the order of 2,000 feet), the CDP airspeed maybe a high value near Vy. This will allow a higher takeoff weight and demonstrate compliance with the V~oss climb requirement of $ 29.67(a)(l). 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 92 Par 60

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requirement of $j 29.67(a)(l). 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 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 69 of this advisory circular.

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 72, a minimum 5-knot clearance from any HV limiting condition should be provided throughout the takeoff flight path (see Figure 64-1 ), 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 l-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 FWVAUTHORITY 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.

Par 60 93

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(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 at CDP: an airspeed equal to or greater than V~oss, 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 V~oss, 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 V~oss, the takeoff distance then becomes the distance to the point in the takeoff profile at which both V~oss 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 V~oss. 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 FANAUTHORITY 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).

94 Par 60

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(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 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 V~oss until approximately 100 feet above the takeoff surface. It should be demonstrated that the rotorcraft including any configuration changes can be accelerated from VTOSSto 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 (HJ. At a given H~, 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 72 of this AC. See Paragraph 55 of this AC regarding altitude extrapolation of test results.

(1O) Ambient Con ditionq. Appropriate test limits for ambient conditions such as wind and temperature are contained in Paragraph 55 of this AC. 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 Par 60 95

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Paragraph 55(b)(l) of this AC 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.

(i) Gener~. 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) Takeo ff Profile. A typical vertical takeoff profile for a ground level heliport is shown in Figure 60-1. 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 60-2. Considerations are similar to those of the ground level heliporl in Figure 60-1; 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 Dec ision Point (C DP). 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.

96 Par 60

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Normally, a minimum l-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.

(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 Reau rement~.i (A) The OEI takeoff profile should include a climb at V~o~~ 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) trapolatiom Basic guidelines for extrapolation are contained in Paragraph 55 of this advisory circular. 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) Niaht Operaticms.

(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.

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(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.

98 Par 60

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All Engine Climbout Accelerate to VY when clear of obstacles CDP + Vertical t I Rejected Climb Takeoff Path FIGURE 60-1. CATEGORY A VERTICAL TAKEOFF PROFILE GROUND LEVEL HELl PORT Par 60 99

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All Engine Climbout i n \ Vertical Climb ++ 7 & Minimum Rejected + 15 Feet Takeoff Path Minimum + 35 Feet FIGURE 60-2. CATEGORY A VERTICAL TAKEOFF PROFILE PINNACLE 100 Par 60

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7130197 AC 29-26 60A. ~ 29.59.29,60.79.61 and 79,67 (Amendment 29-39) TAKE OFF PATH} N~EOFF: GROUJ,4AND El EVATED HELlPORT : CATEGORY A (For $29.59 prior to Amendment 39, see Paragraph 60.)

lanatio~ Amendment 29-39 moved the rejected takeoff requirements from new $29.62 and clearly defined the takeoff path. It also added new ~$j 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 Poinf. 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 (V~o~~). The engine failure point (EFP) and TDP are separated by pilot recognition time. (See Conventional Takeoff Profile, Figure 58A-1, Paragraph 58A of this advisory circular.) TDP is a “go/no-go condition which is analogous to VI 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.

Par 60A 101

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(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 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 V~o~s 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 VToSsand attain an altitude of 35 feet and positive rate of climb and then climb to 200 feet above the takeoff sutiace 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 VTo~s, 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)(l) 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 lA-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 %-minute time limit.

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

(5) Continued Climbou t Path. Continued acceleration and climb capability are assured by the 100 FPM V~oss climb requirement of ~ 29.67(a)(l) and the 150 FPM 102 Par 60A

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requirement of 5 29.67(a)(2), normally demonstrated at VY. It should be shown that the rotorcraft can be accelerated from VToss to Vy in a continuous maneuver without losing altitude, including any nonfigurative change (landing gear retraction, etc.). The distance 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) Establishin a 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 VTos~ climb requirement of ~ 29.67(a)(l). 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 S 29.67(a)(l). 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:

Par 60A 103

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(A) The fiightcrew 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 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) Rejec ted Takeoff Distance. The rejected takeoff is similar in many respects to the height-velocity (HV) tests described in Paragraph 69 of this advisory circular.

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 72, a minimum 5-knot clearance from any HV limiting condition should be provided throughout the takeoff flight path (see Figure 64-1 ), and tests should be conducted simulating an unplanned engine cut. The HV diagram appropriate to the Category A test weights maybe much less restrictive than that determined for Category B conditions. Normally, a minimum l-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 V~o~s, a positive rate of climb, and a height of at least 35 feet above the takeoff sutiace.

If the rotorcraft descends below 35 feet above the takeoff surface while accelerating to V~o~~, the takeoff distance is extended to the point that 135feet 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 V~o~~, the takeoff distance then becomes 104 Par 6oA

### Page 96

the distance to the point in the takeoff profile at which both V~o~~ 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 V~o~~ 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 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 FANAUTHORITY 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 V~o~~ 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, Par 6oA 105

### Page 97

etc.) used to show compliance with the climb requirements of S 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 (H~). At a given H~, 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 Loadinq. 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 due to forward/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 accountability.

(9) Extrapo Iation. Takeoff and landing data may be extrapolated up to 4000 feet along an established W/O 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/a limit. See Paragraph 71 b(5) for further discussion of landing gear loads.

(1O) Ambient Co nditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in Paragraph 55 of this AC. 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 765(a) (3)(iii) of this AC 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.

106 (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 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 60A-I. 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 60A-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 V~os~ with a positive rate of climb is given in the performance chapter of the RFM.

(iii) Takeo ff 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 vettical 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 l-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.

(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 Par 60A 107

### Page 99

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 V~o~~ with a ‘positive rate of climb; for an elevated heliport, the lowest point of the takeoff profile and not less than V~o~~ 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.

(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 V~oss to 200 feet above the takeoff surface then an acceleration in level flight from V~oss 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) maybe met at referenced points located respectively at 200 feet and 1000 feet above the takeoff surface. The distance required to accelerate from V~o~s 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 VTo~~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) maybe 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 55 of this advisory circular. Weight can not be extrapolated above test weight. Altitude extrapolation should be limited to a maximum of* 4000 feet.

(12) Niaht Op erations.

Par 6oA

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7130197 AC 29-2B

(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 recommended takeoff profile. Night OEI rejected takeoffs and continued takeoffs from the TDP should be conducted to ensure adequate night field of view and realization of Category A field lengths.

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

Par 60A 109

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7/30/97

**ALLENGINE**

CLIMBOUT *

**ACCELERATETOVY**

**WHENCLEAROFOBSTACLES**

**OEICLIMBOUT TDP**

VERTICAL CLIMB FIGURE 60A-I . CATEGORY A VERTICAL TAKEOFF PROFILE GROUND LEVEL HELIPORT 110 Par 60A

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7130197 AC 29-2B

**ALL ENGINE**

CLIMBOUT ‘Dp-’x/~ ACCELERATETOV~O=~

VERTICAL CLIMB +&

**REJECTED**

TAKEOFFPATH OEIFLIGHT PATH I [ vTo~~ and positive 1 rate of I climb ~ Takeoff distance ‘~ FIGURE 60A-2. CATEGORY A VERTICAL TAKEOFF PROFILE ELEVATED HELlPORT Par 6oA 111

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AC 29-2B

**61.-63. RESERVED.**

112 (thru 114) 7/30/97 Intentionally Left Blank Par 61

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7/30/97

64. 6 29.63( Amendment. 29-1 2) TAKEOFF: CATEGORY B, AC 29-2B

a. Exdanation.

(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:

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

(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 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 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 (V~Ospeed) 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. Pm edures.

(1) jnstrumentti. A ground station will measure ambient temperature, humidity (if applicable), and wind. For allowable wind conditions and engine power considerations refer to Paragraph 55 of this advisory circular. A photo panel or hand Par 64 115

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recording method 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 57 of this advisory circular. “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 V50 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 55 of this advisory circular.

(4) Test Techniques. For the FAWAUTHORITY 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~o 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).

116 Par 64

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LLl -1 Eatn IA ;

Y s AVOID 5 KTS MINIMUM MARGIN Par 64

**VELOCITY (KNOTS) FIGURE 64-1. CONVENTIONAL TAKEOFF PROFILE CATEGORYB 117**

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AC 29-2B

**65. RESERVED.**

7/30/97 Intentionally Left Blank l% 65

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66. ~ 29.6 5 (Amendment 29 -15) CLIMB: (A LL ENGINES OPERATING).

a. Ex~lanation.

(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 V~E 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 VN~ 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. Procedu e to Deter rmine 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. VVindshear 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 Par 66 119

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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/0). The test is normally started at the desired W/a with maximum continuous power, or at V~~, 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/o. 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 maybe 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 Enaine O~eratina 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.

Par 66

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66A. $29 ,64 and 29.65 (Amen dment 29-39) C LIMB: [GENERAL AND ALL ENGINES OPFRAT ING).

a. Explanat ion.

(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 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 VN~ 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 VN~ 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. Produ e to Deter rmine 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.

par 66A 121

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AC 29-29 7130/97

(2) Level flight petiormance (speed power) may also be used to determine best rate-of-climb airspeed (VY). The testing should be done in smooth air. The 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 (Whs). The test is normally started at the desired W/o with maximum continuos power, or at V~~, 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/O. 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. P~ in limb Perfor a CQ.

(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 maybe 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, 122 Par 66A

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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.

67. 629,67 (Amendment 29-34) CLIMB: ONE ENGINE INOPERATIVE.

a. Exdanation.

(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 (V~o~~) 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 instaIled power available values. The landing gear is retracted and the airspeed is that selected by the applicant.

b. Procedu res.

(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (Paragraph 66) except that the V~o~~ and the Category A climb speed maybe 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 multiengined 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), Par 66A 123

### Page 113

rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb petiormance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilotlcrew 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.

(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 V~o~~ that will result in 100 feet per minute rate of climb at the maximum weight determined in

(b)(4)(i).

67A. $79.67 (Amendment 79-39) CLIMB: ONF ENGINE INOPERATIVE.

a. planat ioq.

(1) Amendment 29-39 expanded the OEI rate of climb requirements. The guidance material presented in Paragraph 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 400 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 beat 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 petiormance must be 124 Par 67

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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 66) except that the V~os~ and the Category A climb speed maybe 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 VTOSSat 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 multiengined 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 V~oss that will result in 100 feet per minute rate of climb at the maximum weight determined in b(4)i.

Par 67A 125

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68. & 29.71 (Amendment 29-12) ROTORCRAFT ANGLE OF GLIDE: CATEGORY B.

(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 engine failure conditions at higher altitudes and the pilot is required to petform 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, maybe 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 68-1.

(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 maybe conducted in conjunction with the climb performance tests. The required data are similar for both tests and it is sometimes convenient and eflicient to run alternating climbs and descents through a 126 Par 68

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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**

AIRSPEED FIGURE 68-1. AUTOROTATIONAL CHARACTERISTICS - TYPICAL ANGLE \JRECOMMENDED /AUTOROTATIONAL SPEED‘k / I/’ \

**MINIMUM BEST GLIDE**

RATE OF DESCENT / / / / / / / / 128 Par 68

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7130197 AC 29-2B

69. 629.73 (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 56).

a. Exdanation.

(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 (lGE). 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 % 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 57 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 Par 69 ’129

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height above the ground. During tethered hover tests, the rotorcraft should beat 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 C#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 maybe 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)

(ii)

(iii)

(iv)

(v)

(vi)

(vii) Engine torques.

Rotor speed.

Ambient temperatures.

Pressure altitude.

Fuel used (or remaining).

Load cell reading.

Generator(s) load.

130 Par 69

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As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in Paragraph 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 FAWAUTHORITY, 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.

70. Q 29,75 [Amendment 29-17) I ANDING.

a. Exdanation.

(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 Par 69 131

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AC 29-2B 7/30197 extrapolation methods are, of course, allowed. General guidance on those subjects is given in Paragraph 55 of this advisory circular. As in other performance areas, engines must be operated within approved limits. Power considerations are the same as those described under Paragraph b(2)(iij(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) ~. 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 70-1.

(A) U. 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 V~o~~ 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 132 Par 70

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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 for the Category A approach condition in comparison to that determined under high power conditions in S 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) /411ena e oin ut 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 80a(2) (iii)(A) of this advisory circular 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 VNEengine failure testing where instrumentation and critical loading have been established for those test conditions. See discussion under Paragraph 80 of this circular. 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) hstrumentat ion/Eaui~m ent. 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 Par 70 133

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AC 29-2B 7/30197 record of rotor RPM at touchdown is necessary to assure it does not exceed transient limits. Rotor RPM at touchdown maybe 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.

(B) Estab Iishing 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)(l). In this case, the requirements ot ~ 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 Ianding from the lower speed or higher altitude decision point for shorter landing distances. In this case the requirements of S 29.67(a)(1) maybe limiting. The climb performance and climb speeds required by ~ 29.67(a)(l) and (2) should be established prior to Category A Ianding tests.

(2) The one engine inoperative landing is similar in many respects to the height-velocity tests described in Paragraph 72 of this advisory circular. 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 l-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 FAWAUTHORITY 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 68 of this advisory circular.

(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 134 Par 70

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control in the cockpit so that engine topping can be accurately adjusted for varying ambient conditions. With such a device in the control system it becomes vitally important to check topping power prior to each test sequence.

(D) Aircraft Loa dinq. 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.

(E) Extrapolation. Weight cannot be extrapolated above test weight.

See discussion under Height-Velocity Testing in Paragraph 72 of this advisory circular.

If no marginal areas are apparent and an acceptable analytical method is used, petiormance data may be extrapolated k4,000 feet density altitude from test conditions.

(See Paragraph 55 of this circular.)

(F) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in Paragraph 55 of this advisory circular. 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 55 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, S 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 80 of this advisory circular. 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/a (gross weightidensity 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 lightweight. This testis 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) Par 70 135

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AC 29-2B 7130197 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 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 sutiace 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 72 of this advisofy circular.) 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 Land inas. The reader should be familiar with the preceding discussion of conventional Category A landing profiles because duplicate information is not repeated here. Atypical vertical landing profile is shown in Figure 70-2. 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 them 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 l-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 136 Par 70

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7130/97 AC 29-2B heliports. The LDP must be clearly defined and flight manual instructions should carefully explain any pilot procedures. An illustration similar to Figure 70-2 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. Catea orv B Reauirements.

(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 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 58 of this advisory circular.

(2) Proce dures. 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 Par 70 137

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first reaches 50 feet above the landing surface to the point at the foremost part of the rotorcraft (including rotor tip path) afler 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 FIVVAUTHORITY 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.

138 Par 70

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vTOSS Vy t50 FT. \ 4 \ \ \\ 4 LANDING DISTANCE b FIGURE 70-1. CATEGORY A CONVENTIONAL LANDING - CLEAR HELIPORT Par 70 139

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v VyTOSS vTOSS \ n b t ) 35 ft.

25 ft. Landing Distance - + 4 140 FIGURE 70-2. CATEGORY A VERTICAL LANDING Par 70

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70A. S6 79.75, 29.77.29,79.29. 81, and 29,83 (Amendment 29-39) LA NDING.

(For ~ 29.77 and ~ 29.79 prior to Amendment 39, see Paragraphs 71 and 72 respectively.)

a. Ianat ion.

(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 70 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, $529.51 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 55 of this advisory circular. As in other performance areas, engines should be operated within approved limits. Power considerations are the same as those described under Paragraph be.

(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 Rea uirements.

Par 70A 141

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AC 29-2B 7130197

(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/after LDP. See Figure 70A-I.

(i) ~. 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).

(ii) Landing distance. Approach and landing path requirements are stated in 5$ 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 142 Par 70A

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7130/97 AC 29-2B 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 Ianding 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 Enaine 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 sutface after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously. See Paragraph 80a(2) (iii)(A) of this advisory circular 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 80 of this circular. 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) Jnstrumentat ion/Eauipment. 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.

Par 70A

(ii) Establishing the LDP.

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AC 29-26 7130/97 (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)(l). 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 72 of this advisory circular. 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 such that recognition is at the LDP with a l-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 71 of this advisory circular.

(iii) Power. Power should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of engine topping to minimum specification values for the range of atmospheric variables to be approved.

This is frequently done by installing an adjustable device in the throttle linkage with a control in the cockpit so that engine topping can be accurately adjusted for varying ambient conditions. With such a device in the control system, it becomes vitally important to check topping power prior to each test sequence.

(iv) Aircraft Load inq. 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 144 Par 70A

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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.

(v) rapolatlon. Landing data may be extrapolated along an established W/a 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 Wia 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 55 of this circular.)

(vi) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in Paragraph 55 of this advisory circular.

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 765 details allowable wind credit.

(vii) All Enaine Out Landinq.

(A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion of this paragraph, ~~ 29.79 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/a 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 multiengined rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to this relatively low inertia level, considerable collective may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which may lead to rotorcraft damage in testing this requirement (e.g., droop stop breakage or loss Par 70A 145

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AC 29-2B 7130197 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 multiengined rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (see Paragraph 72). 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 70A-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 AEO approach procedures are presented in the Flight Manual according to Paragraph 70b(l)(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 146 Par 70A

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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 l-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 fotward-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 70A-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 Reau irements.

(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 Paragraphs 55 and 58 of this advisory circular.

(2) Procedures,

(i) Landina 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 sutface to the point at the foremost part of the rotorcraft (including rotor tip path) alter 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.

Par 70A

(ii) A1l-Enaine-Out Land ing.

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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 GWia 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 multiengined 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 multiengined rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (see Paragraph 72). 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 148 Par 70A

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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 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 limits. 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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vTo=~ Vy v t TOSS 50 ft. + \ i ‘\\ \ ~ LANDING DISTANCE + FIGURE 70A-I . CATEGORY A CONVENTIONAL LANDING - CLEAR HELlPORT 150 Par 70A

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7/30/97 AC 29-2B FIGURE 70A-2. CATEGORY A VERTICAL LANDING par 70A 151

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71. 6 7 Am~ n RY A (For ~ 29.77 after Amendment 38, see Paragraph 70A)

a. Exp Ianation. This rule has two distinct portions.

(1) Section 29.77(a) states that the rotorcraft must be capable of transitioning smoothly from each approved Catego~ 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@ 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 70-1 of this advisory circular.) 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) Instrument ation. Instrumentation requirements are similar to those for Categoty A takeoff. A ground station with positioning capability is needed along with on-board instrumentation of engine and flight parameters.

(2) Balked Landing Prefiles. 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 maybe 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 qualities 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 Paragraph 66 and 67 of this advisory circular.

Par 71

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71A. ~ 29.85 (Amendment 29-39) BALKED LANDING: CATEGORY A (For Balked Landing prior to Amendment 39, see S 29.77 and Paragraph 71.)

a. Explanation. Amendment 29-39 revised and relocated the original ~ 29.77 as a new ~ 29.85. The guidance material of Paragraph 71 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 70A-I of this advisory circular.) The safe climb conditions 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. Instrumentation 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 Landina 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 beforeiafter 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 V~os~ 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.

Par 71A , 153

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(3) Handling Qualities. Handling qualities 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)(l ) and (2) must also be met in the event a balked landing is made. See Paragraphs 66 and 67 of this advisory circular.

(For $29,79 after Amendment 29-38, see Paragraph 70A)

a. ~planatio~.

(1) The height-speed envelope is normally referred to as the height-velocity (HV) diagram. It defines an envelope of airspeed and height above the ground from which a safe power-off or OEI landing cannot be made. The diagram normally consists of three portions: (a) the level flight (cruise) portion, (b) the takeoff portion, and (c) the high speed portion. See Figure 72-1. The high speed portion is omitted on occasions when it can be shown that the rotorcraft can suffer an engine failure at low altitude and high speed (up to V~) 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, Paragraphs 58, 64, and 70 of this advisory circular.

(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 oft 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)). Altitudes below the knee represent takeoff profile points. For test points in the takeoff portion, takeoff power (or a lower power selected by the applicant as an operating limitation), and normal pilot reaction time will be used.

(5) Since the HV diagram may represent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer’s buildup program should be reviewed to determine the amount of conservatism in the HV diagram (if any). It should be remembered that the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup 154 Par 71A

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testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points are approached from above and below respectively. Portions near the knee are initially evaluated at high speed with subsequent backing down of the speed. In most rotorcraft the effective flare airspeed is critical. At airspeeds slightly below this value, the ability to arrest and control descent rates through use of an aft cyclic flare may be greatly diminished, Extreme care should be exercised when “backing down” to lower speeds.

(6) In addition to the on-board and ground instrumentation, a motion picture camera or other position measuring equipment should cover each run.

(7) For FANAUTHORITY 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 FAWAUTHORITY approval. In lieu of pure analytical methods, simulations have been used successfully, especially for multiengined rotorcraft.

In either case, the maximum allowable extrapolation should be limited to 2,000 feet density altitude (H~). HV test weights should be consistent with the takeoff and landing WAT (weight, altitude, temperature) limit curve which will be placed in the Rotorcraft Flight Manual. 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/a. See Figure 72-2, 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 72-2, Part B. Another method of presentation is to show varying weights at a constant density altitude. (See Figure 72-2, Part C.)

(9) 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, Ianding, and HV tests.

(1O) Rotorcraft certificated prior to Amendment 29-21 were required to have the resulting height-velocity 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 height-velocity 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) cetilficated in Category B. Testing must be completed with the aircraft at the maximum gross weight at sea level. For altitudes above sea level, the 155

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test aircraft must be at a weight no less than the highest weight the rotorcraft can hover out-of-ground-effect (OGE). Rotorcraft certificated prior to Amendment 29-21 can update their certification basis to take advantage of this provision.

b. procedures.

(1) jnstrumotation.

(i) Ground Stat io~, 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) ~~ Rotorcraft). Necessary installed test equipment may include photo panels and/or recorders for recording engine parameters, control positions, landing gear loads, landing gear deflections, airspeed, altitude, and other variables. An external light attached to the rotorcraft (or any other means of identifying the engine failure point to the ground camera or phototheodolite) is needed to identify the exact time of engine failure and may also be used to synchronize the ground recorder with the airborne recorded data.

(2) Analytical Prediction. The HV diagram can be estimated by analytical means and this is recommended prior to test. HV, however, is the least predictable of all rotorcraft performance and because of this, the expansion and extrapolation of test data must be done with great care. Test weight may not be extrapolated. All test points should be approached conservatively with some speed or altitude margin. If the 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 andior simulation method must be utilized for extrapolation.

(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 156 Par 72

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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 normal Category B rotorcraft. For Category A rotorcraft, the desired topping power (for the remaining engine(s)) should be set prior to the test. This power value will need adjustment as ambient conditions change. The power can be takeoff power (TOP), 2 lf2-minute power, or some calculated lower power for simulating hot day or higher density altitude conditions. Power is verified and recorded by the pilot by “topping” the engine(s) prior to engine failure tests. Care must be taken to 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 Loadin~. Weight extrapolation is not permitted for HV. Therefore, the test weight must be closely controlled. Ballast or fuel should be added frequently to maintain the weight within -1 to +5 percent when testing final points. Ordinarily tests are conducted at a mid center of gravity unless a particular loading is expected to be particularly critical.

(5) Landina 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 Con siderations. 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) Hiah Hove r. A stabilized out-of-ground-effect (OGE) hover condition prior to power failure is essential. A minimum l-second time delay between power failure and initial control actuation is utilized. Following the time delay, the primary concern is to quickly lower collective and to gain sufficient airspeed to allow an effective flare approaching touchdown. While the immediate development of airspeed is necessary, the dive angle must be reasonable and must be representative of that expected in service. While initial aircraft attitude will vary between models and with changing conditions, 10°-200 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 Par 72 157

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stabilized OGE hover condition prior to power failure is essential. If a stabilized high hover condition cannot be achieved prior to the engine cut, then this point should be tested from a minimum level flight speed. This will result in an open-ended HV diagram. A smoke source or balloon on a long cord is highly desirable since the wind can vary significantly from surface observations to typical high hover altitudes. Vertical speed must be very near zero at the throttle chop. Any climb or sink rate can have a significant influence on the success of the test point. Use of a radar altimeter with a cross check to barometric altitude is essential,

(ii) Low Hover. From the low hover position there is no flare capability and little time for collective reaction. No time delay is applied other than normal pilot reaction. For typical designs the collective may not be lowered after power failure.

Lowering of the collective is not permitted because it is not a pilot action which could be expected if an engine failed without notice during a hovering condition in service. Initial lowering of collective immediately after power failure can result in very high, unconservative low hover altitudes that are unrealistic for operational conditions. If, however, a design is such that a l-second pilot delay after power failure could be achieved without any appreciable descent, a slight lowering of collective could be allowed.

(iii) Takeoff Corridor. Normal pilot reaction is applied when the engine is made inoperative. At low speeds collective may be lowered quickly to retain RPM and minimize the time between power failure and ground contact. If airspeed is sufficient for an effective flare, the aircraft is flared to reduce airspeed, retain rotor RPM, and control vertical speed prior to touchdown. Considerable surface area may be needed for a sliding or rolling stop.

(iv) Additional Considerations. The “in-between” points utilize similar techniques. The cruise or “fly-in” points are similar to the high hover point although the steep initial pitch attitudes are not needed as altitude is decreased and airspeed is increased along the curve. The low speed points along the takeoff corridor are similar to the low hover point except that the collective may be quickly lowered and some flare capability may be used as the “knee” is approached. The pilot should be proficient in all normal autorotation landings before conducting HV tests in a single-engine rotorcraft.

(7) ~!l. 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 heightkpeed 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, 158 Par 72

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7130/97 AC 29-2B winds above permissible levels, rotor droop below approved minimum transient RPM, damage to the rotorcraft, excessive power, incorrect time delay, etc., cannot be accepted.

(ii) After the HV diagram is defined, it should be ascertained that the corridor permits takeoffs within *5 knots of the recommended takeoff profile.

(9) Flight Manual. The flight manual should list any procedures which may apply to specific points (e.g., high speed points) and test conditions, such as runway surface, wave height for amphibious tests, marginal areas of controllability or landing gear response, etc. The HV curve should be presented in the RFM using actual altitude above ground level and indicated airspeed.

(1O) Night Evaluation. If a rotorcraft is to be certified for night operation, a night evaluation is required. Engine failures should be conducted along the recommended takeoff path. Landings should also be qualitatively evaluated with an engine failed. Engine failures at critical HV conditions are not required. The intent is to show adequate visibility using aircraft andlor runway lights without requiring a duplication of the daytime HV test program. See related discussion under Paragraph 64 of this advisory circular.

(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) \ LEVEL FLIGHT r TAKEOFFCORRIDORhTAKEOFF

SPEED (V) FIGURE 72-1. HEIGHT -VELOCIW DIAGRAM 160 Par 72 ,..,,.. .

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Height H~ f ‘e[”city l-~d \ c Approx.

Constant GW/6 1 Weight

CONSTANT HV DIAGRAM, VARIABLE WEIGHT Height Height Increasing Density Altitude I Velocity LB J-

CONSTANT WEIGHT Increasing Gross Weight Velocity L_.- c__l

CONSTANT DENSITY ALTITUDE FIGURE 72-2. ALTITUDEANEIGHT ACCOUNTABILITY Par 72 161

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(For Limiting Height-Speed Envelope prior to Amendment 39, see ~ 29.79 and Paragraph 72.)

a. Exp anatI ion. Amendment 39 redesignated $29.79 as ~ 29.87.

b. Procedures. The guidance material presented in Paragraph 72 continues to apply.

**73.-78. RESERVED.**

162 (thru 168)

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**7/30/97 SECTION 4. FLIGHT CHARACTERISTICS AC 29-2B**

79. 29.1 44 (Amendment 29-24) FLIGHT C HARACTERISTIC S - GENERA4.

a. ~planation.

(1) This regulation prescribes the general flight characteristics required for certification of a transport category rotorcraft. Specifically, it states that the rotorcraft shall comply with the flight characteristics requirements at all approved operating altitudes, gross weights, center of gravity locations, airspeeds, power, and rotor speed conditions for which certification is requested. The reference to “altitude” in $ 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 Paragraph 775 of this advisory circular.

(2) Generally the aircraft structural (load level) survey accounts for takeoff power values at speeds up to and including VY. At speeds above VY, maximum continuous power is assumed. Stress to rotating components usually increases with airspeed and power. If the takeoff power rating exceeds the maximum continuous power rating, and the structural survey has been conducted under the assumption that takeoff power is not used at speeds above VY, the Rotorcraft Flight Manual must limit takeoff power to speeds of VY and below. If takeoff power is structurally substantiated throughout the flight envelope, and appropriate portions of the controllability, maneuverability, and trim requirements of 5$29.141 through 29.161 are met at takeoff power levels, no flight manual entry is needed. Obviously, if transmission limits for 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.

Par 79 169

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AC 29-2B 7/30197

(4) During the FANAUTHORITY 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 opinion and to carefully document the results of these evaluations.

Section 29.141 (b) provides the regulatory basis for these strength and skill requirements. The general requirements for a smooth transition capability between appropriate flight conditions are also included in $j 29.141(b). These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft,

(5) For night or IFR approval, ~ 29.141(c) contains the general regulatory reference which requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this order.

80. ~ 29.143 (Amendme nt 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. 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), $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 V~~ with critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at V~~ 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 nosedown 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 V~~ 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.

170 Par 79

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(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, 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 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 Section 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 FAR 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 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(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, Par 80 171

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and high speed flight up to VNE. Entry conditions for the first engine failure are engine or transmission limiting maximum continuous power (or takeoff power where appropriate) including reasonable engine torque splits. For multiengined Category A installations (three or more engines) subsequent engine failures should be conducted utilizing the same criteria as that used for first-engine failure. The applicant may limit his flight envelope for subsequent failures. Initial or sequential engine failure tests are ordinarily much less severe than the “last” engine failure test required by ~ 29.75(b)(5).

The conditions for last-engine failure are maximum continuous power, or 30-minute power if that rating is approved, level flight, and sudden engine failure with the same pilot delay of 1 second or normal pilot reaction time, whichever is greater.

(B) For 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 VNE (power-on) and conditions of hover, takeoff and climb at VY. Maximum continuous power is specified prior to the failure for the cruise condition. Power levels appropriate to the maneuver should be used for other conditions. The corrective action time delay for the cruise failure should be 1 second or normal pilot reaction time (whichever is greater). Cyclic and directional control motions which are 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 l-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 maximum continuous (MC) power, it does not limit engine failure testing to MC power. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions 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 VN~ (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 V~~ (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 ~ 29.1 505(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 172 Par 80

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testing at V~~. Sufficient directional control must exist to allow straight flight in autorotation during all approved maneuvers including 30” banked turns up to V~~ (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 VNEat MC power per ~ 29.143(b), 1.1 VNE at power for 0.9 V~ per $ 29.175(b) or $29.1505, and to 1.1 V~E (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 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 fotward and lateral cyclic available to reach the power-on VN~. 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.

(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 Ieftward 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 Par 80 173

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reduce left sideward flight capability. In the case of fly-by-wire surfaces, both the high speed forward flight controllability and the rearward flight capabilities could be affected.

Flight control systems that incorporate automatic devices should be thoroughly evaluated for critical areas. Every failure condition that is questionable should be flight tested with the appropriate actuator fixed in the critical failure position. These failures may require limitations of the flight envelope. Any procedure or limitation that must be observed to compensate for an actuator hardover and/or softover malfunction should be included in the Rotorcraft Flight Manual.

b. Procedures.

(1) Flight test instrumentation 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 V~~ controllability testing is ordinarily aft CG, MC power, and minimum power-on rotor RPM, although power and RPM variations should be specifically evaluated to verify their effects. The turbine engine is sensitive to ambient temperatures which affect the engine’s ability to produce rated maximum continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit maximum continuous power would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward CG should be spot checked for any “tuck under” tendency at high speed. The VNEcontrollability 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 VNE varies with altitude or temperature, VNEfor existing ambient conditions is utilized for the test. Extremes of the altitude/temperature envelope should be analyzed and investigated by flight test.

(3) The critical condition for 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 174 Par 80

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7130197 AC 29-2B somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the rotorcraft is held in the most critical position relative to wind.

(4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of his engine, drive system, and rotor limits. These limits were established during previous ground and flight tests and 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 V~~ with any power setting must permit suitable nose down pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft’s response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft motion, the results are unacceptable; e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral cyclic to correct roll in conjunction with forward cyclic to correct pitchup is satisfactory.

Obviously during the conduct of these tests, all available techniques should be utilized when the pilot finds himself “out of control.” However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion.

(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all engine inoperative requirements of ~ 29.75(b)(5) and (c). See Paragraph 70 of this advisory circular.

Par 80

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AC 29-2B 7/30197

81.

a. at[om 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 pilot-induced 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 machine. In the case of an irreversible design equipped with an artificial force feel system in pitch and roll, excessive friction can mask a shallow force gradient making positive stick centering and control force static stability difficult if not impossible to demonstrate. In such an instance, the initial choice of fixes might include implementation of a steeper force gradient or addition of a force preload. Unfortunately, these solutions often lead to the kinds of problems discussed earlier. Care must therefore be exercised during the initial design phase to ensure that the components and characteristics of the flight control system are well matched.

b. Procedures. Regardless of the flight control system sophistication, it is important that the test pilot understand the system configuration prior to flight evaluation. Appropriate mechanical characteristics should be documented. For VFR aircraft, the mechanical characteristics are typically assessed in flight on a qualitative basis. If a controllability or workload problem is identified, a more detailed investigation would be necessary. Since IFR certification rules include specific trim and force requirements, a more quantitative investigation of mechanical characteristics is normally conducted. The constantly varying feedback forces of reversible flight control systems generally make such designs unsuitable for IFR application. Irreversible system mechanical characteristics can often be partially documented on the ground with external hydraulic and electrical power supplies connected to the aircraft.

Knowledge of the breakout, friction, and force gradient characteristics prior to flight can be useful to the pilot during flight evaluation of the system.

82. $29.161 TRIM CO NTR04.

a. Expla nation.

(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 reference 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.

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(2) A number of devices maybe 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 Paragraph 775 of this advisory circular.

82A. ~ 29.161 (Amendment 29-24) TRIM CO NTROL,

a. Exdanation. 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.

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83. 29.171 STABILI TY: GENERAL.

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

b. Bxxedws.

(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 FANAUTHORITY 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.

84. s 79.173 (Ame dment 29-24) sn TATIC LONG TUDI INAL STABILITY.

a. Exp lanatio~.

(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 178 Par 83

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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 85 of this advisory circular for an explanation of detailed flight test procedures.

85. S 29.175 (Amendme t 29 -?4) DEMONSTRATION OF STATIC ~.

a. Explanation.

(1) This rule incorporates the specific flight requirements for demonstration of static longitudinal stability. Specific loadings, configurations, power levels, and speed ranges are stated for conditions of climb, cruise, autorotation, and hover.

(2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in $29.173, which states that throttle and collective pitch must be held constant, is appropriate for administration of this rule, and the specified power in ~ 29.175(a), (b), and (c) should be considered as power established at initial trim conditions. This will result in slightly higher or lower torque readings at “off trim” conditions. Collective and throttle controls are held constant when obtaining data during climb, cruise, and autorotation tests.

(3) The effects of rotor RPM on autorotative static stability should be determined, and positive stability demonstrated for the most critical RPM. For Category A rotorcraft this requirement maybe satisfied at a nominal RPM value. RPM values can be expected to change as airspeed is varied from the “trimmed” condition.

Manufacturer’s recommended autorotation airspeed is ordinarily used for trim.

(4) Hovering is considered a flight maneuver for which the pilot repeatedly adjusts collective to maintain an approximately constant altitude above the ground. For hover stability tests, collective and throttle adjustments are made as necessary to Par 84 179

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AC 29-2B 7/30197 maintain an approximately constant height above the ground. Also, a limited amount of negative longitudinal control travel is allowed with changes in speed.

b. Procedures.

(1) Instrumentation.

(i) Sensitive control position instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when using a “before and after” method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it and attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The Figure 85-1 plot is a typical presentation of longitudinal static stability.

(ii) Other necessary parameters include pitch attitude, pressure altitude, ambient temperature, and indicated airspeed (pace vehicle or theodolite speed for hover tests). For hover tests, hover height (radar altitude if available), and surface winds should be documented. Two-way communications with a pace vehicle is highly desirable. Ground safety equipment is desirable.

(2) Ambient Conditions. Smooth air is necessary for stability testing. Allowable wind conditions for hover stability testing are the same as those for hover controllability tests and are described in that section (Paragraph 81). Extrapolation is covered in Paragraph 58 of this advisory circular.

(3) Loading. Aft center of gravity (CG) is ordinarily critical for longitudinal stability testing, although high speed flight and hover should be checked at full forward CG and maximum weight. At aft CG, light or heavy weight conditions can be critical.

The manufacturer’s flight data should be reviewed to determine critical loading conditions.

(4) Conducting The Test.

(i) The rotorcraft should be established in the desired configuration and flight condition (climb, cruise, autorotation) with the required power and rotor speed at the trim airspeed. The collective stick should be fixed in that position, usually by applying sufficient friction to insure that it is not inadvertently moved. For autorotative tests, a rotor speed should be selected so that the variations in rotor speed as airspeed and altitude change do not exceed the allowable limits. This point is recorded as the trim point. Airspeed is then increased or decreased in about 10-knot increments, stabilizing on each speed and recording the data. At least two points on each side of the trim speed should be taken.

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(ii) The cruise test should be conducted by varying airspeed around the desired altitude with throttle and collective fixed. This should be accomplished by first determining V~ (level flight speed at maximum continuous power) at the test altitude.

Then reduce power to establish a level trimmed condition at 0.9 V~ (or 0.9 V~~ if lower).

This point is then recorded as the trim point.

(iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually t2,000 feet), first increasing airspeed as data points are collected, then decreasing speed through the same altitude band. It will probably not be possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit. Generally, it will be possible to acquire all the high speed points on one pass and the low speed points on the second.

(iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise, data can be obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise data because collective and throttle are not moved as airspeed is changed at a constant altitude. A typical sequence of speeds that could produce these results would be: 150 (V~), 135 (0.9V~) trim speed, 125, 145, 115, 155, 105, and 165.

(v) For rotorcraft with high rates of climb, a series of climbs, each at a different speed, may be required through a given altitude, utilizing sensitive instrumentation to assure collective position is the same for each data point. In autorotation, a similar case arises and a series of descents, each at a different speed, may be required through a given altitude band, using sensitive instrumentation to assure a repeatable collective position.

(vi) Hover tests should be conducted by maintaining an approximately constant altitude above the ground at the hover height established for performance purposes. The test altitude above the ground may be increased to provide reasonable ground clearance during rearward flight. Groundspeed is varied using a pace vehicle, theodolite, or other velocity measuring equipment. A pace vehicle is an aid in maintaining an accurate hover height. The pilot can accurately maintain height by controlling his sight picture of the pace vehicle (level with the roof, antenna, etc.).

Hover stability tests are ordinarily conducted in conjunction with hover controllability tests because instrumentation and facilities are essentially the same.

(vii) Normally climb, cruise, and autorotation tests should be conducted at low, medium, and high altitudes. See Paragraph 55 for guidance on interpolation and extrapolation. High speed stability has been critical during cold weather testing. [n two par 85 181

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recent models, VNE at cold temperatures has been limited by the stability requirements of ~ 29.176(b). Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather VN~ to tip Mach number values demonstrated during warm weather testing.

(viii) Hover stability should be verified at low altitude and, if required, at high altitude. Refer to Paragraph 55b(2) for guidance on expansion and extrapolation of altitude.

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o CRUISE ❑ CLIMB

A AUTOROTATION ~ HOVER i=LL < -1 -1 3LL 2 0(YIL - (!)z o-1 100 90 80 70 60 50 40 30 20 10 0

**SHADED POINTS REPRESENT TRIM 0 20 40 60 80 100 120 140 160**

CALIBRATED AIRSPEED (KNOTS) FIGURE 85-1. STATIC LONGITUDINAL STABILITY Par 85 183

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86. ~ 29.177 (Amendme nt 29-24) STATIC DIRECTIONAL STABILITY.

a. hplanat ion. This rule requires that positive static directional stability be demonstrated at the ~ airspeeds defined in $29.175. The trim speed for climb is VY and for cruise is 0.9VH or 0.9VNE (whichever is less). For autorotation that airspeed defined by the midpoint of the speed range specified in ~ 29.175(c) may be used,

b. procedures.

(1) Tests for static directional stability require instrumentation for pedal position and sideslip angle. Lateral cyclic control position instrumentation should be provided for IFR certification tests. To obtain accurate sideslip angle and airspeed information, a “yaw boom” is usually installed for the purpose of mounting a sideslip vane and swiveling airspeed pitot head outside the main rotor downwash region of influence.

Special care should be taken to ensure that the yaw boom installation has been verified to be structurally adequate and free of dynamic instabilities for all combinations of airspeed and rotor speed likely to be experienced during the static directional evaluation. For some installations, the instrumentation yaw boom may influence the flying qualities of the rotorcraft itself. Thus, it is advisable to correlate yaw string displacement or slip indicator ball widths of skid with yaw boom sideslip angle, and then repeat a few critical points with the yaw boom removed.

(2) For some rotor system designs, the main and tail rotor flapping angle may be a critical instrumentation requirement for static directional testing. Both main and tail rotor flapping may increase dramatically at high airspeeds with increasing sideslip angle. Therefore, for rotor systems exhibiting this characteristic, flapping should be monitored carefully during the sideslip maneuver to avoid exceeding limitations. Static directional stability is normally defined in terms of pedal displacement required to maintain a straight flight path sideslip. A single-rotor rotorcraft flying in coordinated flight will exhibit a small inherent sideslip due to tail rotor thrust and fuselage/main rotor sideforces. This condition is normally taken as trim with the inherent sideslip angle noted. Airspeeds should be the trim values described above. A generally accepted technique follows:

(i) Stabilize at the trim point, and note indicated airspeed.

(ii) Record trim conditions including inherent sideslip. Maintain fixed collective and throttle for the remainder of the maneuver.

(iii) Smoothly yaw the aircraft with directional control and coordinate with lateral control to establish the desired sideslip angle. A steady heading can best be ensured by maintaining a track over a straight landmark on the ground such as a section line or straight segment of powerline or highway.

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(iv) Note airspeed immediately upon completion of the yaw maneuver.

There may be a small change from the trim airspeed. Fly the new airspeed while maintaining a constant heading, and record indicated airspeed, control positions (directional at a minimum), sideslip angle, rotor speed, rate of descent, amount of ball deflection, and bank angle. The pilot should note the physical sideforce feel experienced. A minimum of two sideslip data points on each side of the trim point should be obtained to adequately define the slope of the pedal displacement versus sideslip angle relationship.

(v) Smoothly return the aircraft to the inherent sideslip angle. Static directional stability plots can be expected to differ slightly on either side of the inherent sideslip angle. Positive static directional stability is indicated by increased left pedal displacement for a larger right sideslip and, conversely, increased right pedal for a larger left sideslip angle.

87. ~ 29.181 (Amendment 29-24) DYNAMIC STABILITY: CATEGORY A ROTORCRAFT.

a. Explanation. This section requires that Transport Category A rotorcraft, certificated under Amendment 24 of FAR 29, demonstrate positive damping for short-period oscillations (5 seconds or less) at forward speeds from VY to V~~ with the cyclic, collective and directional controls held in the desired test condition or released by the pilot. This requirement would prevent persistent or divergent short-period oscillations and thus alleviate the pilot workload to actively dampen oscillatory motions for all types of operations.

b. Procedures.

(1) Tests for short period dynamic stability are carried out in the same manner as for IFR (reference Paragraph 775 of AC 29-2A) except the oscillation need not be damped as heavily (i.e., to % amplitude in not more than one cycle). Similarly pulses and doublets may be used to generate an upset condition that would be expected to be encountered in moderate turbulence for that particular rotorcraft.

(2) Tests should be conducted at the critical gross weight, altitude, center of gravity, rotor RPM, and power conditions during routine climb, cruise, and descent condition for speeds from Vy to VNE. This test must be conducted with the minimum amount of stability augmentation approved for continued safe flight. Consideration should be given to optional equipment that are to be mounted externally.

(3) This requirement is not applicable to transport category rotorcraft certificated as Category B only. The requirements for this situation are unchanged.

**88.-95. RESEWED.**

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IN.~ H RACTERISTI

96. s 2~G A RA CTERISTICS).

a, ~ ion. The rule states: “The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.” In addition, $529.235, 29.239, and 29.241, contain specific requirements concerning ground and water handling characteristic evaluations.

b. procedures.

(1) During the flight test program and the F&R program ($21 .35( b)(2)), the rotorcraft will be subjected to evaluations at various weight and CG conditions. Any uncontrollable tendencies found during these test programs must be corrected.

(2) Controllable or damped vibrations or oscillations on the ground or in the water are acceptable, provided the design limits of the rotorcraft are not exceeded.

(3) Any significant vibration or oscillation characteristics found during tests should be described in the test report, and the rotorcraft flight manual should contain appropriate descriptions and procedures to describe and either avoid or handle significant characteristics.

(4) For rotorcraft equipped with wheel gear, the evaluation should include takeoff, landing, and taxi at the maximum airspeed and ground speed CG extremes. If a nose or tail wheel lock/swivel control is installed, each position should be evaluated for limiting takeoff, landing, and taxi speeds. Maximum substantiated speed values should be included in the RFM as limitations.

(5) For water operations, the wave height and frequency or “sea state” should be included as a limitation or, if no limit was reached during testing, the demonstrated values should be placed in the Performance Section of the Rotorcraft Flight Manual.

Information or limits on the allowable “sea state” for rotor startup and shutdown should also be included.

97. s 29.235 TA XIING CO NDITION.

Explanation. The rotorcraft is designed for certain landing load factors ($S 2;.471 and 29.473). The rotorcraft must not attain a load factor in excess of the design load factor when taxied over the roughest ground that may reasonably be expected in normal operation at the expected taxi speeds. This rule applies to wheel landing gear equipped rotorcraft.

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b. Procedu res. The structural substantiation data contains the allowable design Iimitsfor the rotorcraft. Acalibrated accelerometer orloadfactor ``g''meter should be installed, as near as practicable to the rotorcraft CG, to record the maximum vertical load factor attained. Instrumentation of the landing gear and/or related structure may also be an acceptable means of showing compliance.

(1) Calibrated instrumentation should be installed to record the maximum loads or maximum vertical load factor attained during the taxi tests.

(2) The taxi surface should be evaluated for compliance with the rule, Corrugated surfaces, as well as broken or uneven surfaces, in accordance with the rule, should be used.

(3) Representative typical taxi speeds, up to the maximum selected by the applicant, should be attained over the selected taxi surfaces.

(4) A light and heavy rotorcraft weight condition should be evaluated.

(5) Limitations appropriate for the rotorcraft design should be included in the flight manual. If these tests indicate that it is unlikely that limit load factors will be attained while taxiing, flight manual limitations may not be necessary.

(6) Pertinent taxi information obtained from these test conditions may be included in normal procedures of the flight manual.

**98. 629.239 SPRAY CHARACTERISTICS.**

a. Explanat ion. The intent of this requirement is to evaluate by demonstration that water spray does not obscure visibility (day or night) or damage the rotorcraft during normal waterborne operation (for those rotorcraft which have waterborne or amphibious capability).

b. Procedures.

(1) The following maneuvers should be evaluated in ambient conditions up to the proposed sea state or wave height for operation.

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Con- Rotor Altifig. Condition Weight CG RPM tude Remarks 1 Taxi Max Optional Max SL Speeds up to maximum proposed for water operation.

2 Hover Max opt Max - Determine critical hover height, any.

3 Takeoff Max opt Max SL Unstick at maximum proposed water operation speed.

4 Land Max opt Max SL Touchdown at maximum proposed for water operation.

5 Shutdown Opt opt - SL Shut down the rotorcraft.

6 Start Max opt Max SL Start engines and release rotor brake.

(2) The maximum sea state or wave height evaluated under this rule should be stated and included in the limitations section of the flight manual.

(3) The effect of saltwater contamination and deterioration of turbine engines and other component parts of the rotorcraft should be considered in accordance with $29.609 and Paragraph 244 of this advisory circular. Information on saltwater effect and attendant corrective action should be provided in the flight manual, if appropriate, and in the maintenance manual.

99. ~ 29.2 41 GROUND RESONANCE.

a. Exdanation.

(1) The rule states: “The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning.” This rule is a flight requirement that pertains to demonstrating freedom from dangerous oscillations on the ground. CAR Part 7, predecessor to FAR Part 29, originally contained a “strength requirement,” under ~ 7.203, requiring ground vibration tests. This test would identify critical vibration frequencies and modes of the rotorcraft. CAR Part 7, Amendment 7-4, effective October 1, 1959, removed this ground vibration requirement because the agency concluded that if any major component has a natural frequency which could be excited by some operating parameter, such a condition would be revealed in the course of other ground and flight tests. The Federal Aviation Administration (FAA) apparently 196 Par 98

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7/30197 AC 29-2B was depending on demonstrations under $7.1 31/$ 29.241 and the flight load survey data (~ 29.571) to satisfy the objective of the vibration test. However, FAR 29, Amendment 29-3, contained new ~ 29.663 adding reliability and damping action investigation requirements for ground resonance prevention means. A ground vibration survey was not reinstituted by the adoption of ~ 29.663. Compliance with $29.663 does require investigation and substantiation as stated. See Paragraph 268 of this document for compliance with $29.663.

(2) “Ground resonance” is a mechanical instability of the aircraft while in contact with the ground, often when partially airborne. Stated another way “ground resonance” is a self-excited mechanical instability that involves coupling between the in-plane motion of the rotor blade and the motion of the rotorcraft as a whole on its landing gear (reference “Aerodynamics of the Helicopter,” Gessow & Myers, page 308). It is caused by the motion of the blade in the plane of rotation (called in-plane vibration) coupled with a rocking or vertical motion of the aircraft as a whole, The tires, landing gear, and rotor restraint pylon structure act as a spring with a vibration frequency which coincides or couples with the natural in-plane frequency of the blade about a real or effective drag hinge in the plane of rotation. When the frequencies of the two motions (rotor and airframe) approach each other and couple, a violent shaking of the aircraft may occur which, if undamped, could result in the destruction of the rotorcraft.

(3) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities. This mode of vibration or resonance can happen in-flight (called air resonance) as well as on the ground and should be addressed in the certification program. The evaluation should include variations in stiffness and damping that could occur in service to the rotor pylon restraints (reference “Ground Vibrations of Helicopters,” M.L. Deutsch, JAS, Vol. 13, No. 5, May 1946). See Paragraph 268 of this document for the investigation of the variations.

(4) Ground resonance may be prevented by placing the first order in-plane vibration frequency above the rotor turning speed.

(5) For such configurations which are not susceptible to ground resonance (first order in-plane frequency above rotor turning speed), a simple rotor RPM run-up and run-down with appropriate cyclic control displacement (i.e., excitation of any inherent vibrations) is adequate demonstration that a ground resonance condition does not exist.

Unhinged “rigid” rotors, such as Bell Helicopter 2 blade designs, are this type of rotor system.

(6) For configurations that are susceptible to ground resonance (i.e., first in-plane frequency is below the rotor turning speed), ground resonance is generally prevented by dampers on the blade, acting in the plane of rotation, dampers on the landing gear (sometimes serving as oleo struts), or proper placement of the landing gear frequencies combined with rotor and/or landing gear dampers.

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AC 29-2B 7/30197

(7) Elastomeric components (in the rotor pylon support system, possibly in the landing gear, and possibly in the rotor head) are significantly affected by ambient temperature prior to warm-up. Their damping characteristics require thorough investigation for the range of rotorcraft operating environment as noted in Paragraph 268 of this document.

b. Procedures.

(1) in operation, the resonance characteristics should be checked during takeoff and landing at zero speed and during run-on landings using various power values. Under all conditions, any oscillations which may be introduced should be damped. However, no instability should occur at any operating condition such as during RPM changes from minimum to maximum and idle to maximum. For rotorcraft with wheel gear, uneven taxi surfaces in conjunction with particular taxi speeds, may excite ground resonance and should be evaluated by taxiing on typical surfaces, This evaluation may be conducted in conjunction with tests of ~ 29.235.

(2) Slow vertical landings for each configuration are made to establish the touchdown collective pitch angle for each rotor speed. For those aircraft equipped with Stability Augmentation Systems (SAS), all ground resonance investigations should be conducted with SAS on and SAS off. This includes the hovering and running takeoffs and landings, taxi tests, and specific ground resonance tests noted herein.

Consideration should be given to conducting tests in various SAS configurations such as roll channel on, pitch channel off, where such configurations are possible and authorized.

(3) For each rotorcraft configuration tested, the aircraft should be positioned on the ground in flat pitch with the rotor stabilized at the minimum practical rotational speed, or optionally, at a speed shown analytically to have significant margin from indicated resonant conditions. Control system inputs should be used to disturb the system for evaluation of subsequent damping.

(4) For each incremental increase in rotor speed and for each rotor speed setting at increments of collective pitch settings, cyclic and collective inputs should be investigated prior to proceeding to the next rotor speed setting. These inputs should cover the appropriate range and combinations of amplitude and frequency.

(5) Cyclic pitch inputs should be made, either by the pilot through the cyclic stick, or through a signal generating device working in conjunction with the cyclic controls. For each frequency of input, amplitude of the inputs should be increased incrementally and ultimately should be large enough to generate responses representative of normal ground and flight operation on the rotor and support system.

The inputs should continue for a time sufficient to execute five complete counterclockwise circles of the cyclic stick (about neutral) at the selected frequency.

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(6) At each amplitude of cyclic input, the excitation frequency should be incrementally increased over the range of the blade in-plane frequency in the fixed system. Rotor speed settings should be increased to 1.05 times the maximum power-on rotor speed. Collective pitch settings should be increased in increments of not more than 20 percent to maximum collective or alternately to the collective setting required to become partially airborne (when the cyclic is displaced as noted).

(7) Typically, articulated rotor aircraft have natural frequencies on the blade in lag of approximately 0.3 times the power-on main rotor RPM; soft in-plane rotors have natural frequencies approximately 0.7 times the main rotor RPM. Therefore, for example, for a rotorcraft with an in-plane frequency of 0.3/rev, operating at 300 RPM, and with 6 inches of total lateral cyclic stick displacement, the stick should be rotated for 5 revolutions in a 0.6-inch diameter circle at ((1-.03) x 300 RPM) or 3.5 cycles per second to attempt excitation of possible resonant frequencies. At the conclusion of the excitation, the cyclic stick should be returned to the neutral position while continuing the recording of data listed in Paragraph b(l 3).

(8) The complete program should again be repeated with cyclic excitation inputs from the directional and longitudinal controls, if critical for the type of rotorcraft being evaluated.

(9) If onset of ground resonance is encountered, the typical recommended corrective action is to increase the collective pitch and rotor speed and become airborne. However, lowering the collective pitch has been effective for some designs and is considered a satisfactory procedure if resonance can be consistently avoided.

(1O) Landings should be made at the maximum touchdown speed proposed with the rotor speed stabilized.

(11 ) Special Considerations:

(i) The influence of variables including environmental effects, corresponding aircraft component characteristic changes, operational parameters, and surface conditions should be investigated over the ranges proposed for certification.

Additionally, the potential of misservicing and possible failure modes should be evaluated. For ground resonance qualification, where practical, variations from the baseline test configuration may be accomplished by either ground run ($ 29.663(b)) requires investigation of probable ranges of damping), analyses, component tests, aircraft shake test, the specification of special operational procedures in the rotorcraft flight manual, or combination thereof. Detailed and rational analyses showing acceptable correlation to the baseline tests, and for which the input parameters were verified by drawings, calculations, component static or dynamic tests, or by aircraft shake tests simulating the conditions/configurations in question, may be used to limit testing to only those variables and operational conditions showing marginal or Par 99 199

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unacceptable system damping. All operational limitations should be clearly stated in the rotorcraft flight manual. A report of the analytical and/or test results should be permitted per ~ 29.663.

(ii) Potential instability while airborne, called “air resonance” may occur due to the dynamic coupling of the rotor flexibility and the pylon restraint flexibility. The same considerations apply to air resonance as to ground resonance except that the pylon restraint variables replace the landing gear variables. Air resonance should be addressed in the certification program.

(iii) When operating on the ground, there maybe a tendency for the aircraft to exhibit a “ground bounce.” For many configurations, this is a benign, although undesirable phenomenon which may be aggravated by pilot induced oscillations (PIO), particularly if there is little or no friction on the collective.

(12) On rotorcraft with fully articulated rotor heads equipped with landing gear oleos in either skid or wheel configuration, there are tendencies for ground bounce to occur when light on the oleos, either just prior to takeoff or just after landing contact, or during a power assurance check. This bounce may induce ground resonance, particularly if the intensity of the bounce is aggravated by PIO. The corrective action is either to lift off to a hover or to positively lower the collective and remain on the ground.

(13) Instrumentation and Data Acquisition.

(i) Atmospheric Conditions (to be manually noted):

Altitude OAT Wind Velocity

(ii) Aircraft Configuration (to be manually noted):

200 Gross Weight C.G.

Tire Pressure Landing Gear Oleo Pressure Par 99

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(iii) Instrumentation (for recording during test).

Main Rotor RPM.

Time history of cyclic control fore-and-aft and lateral stick position Time history of collective control stick position Time history of rotor damper motion* Time history of pylon component motion* Time history of landing gear (oleo) motion* Time history of aircraft motions* *As required to obtain modal damping

100.-109. ~.

Par 99 201 (thru 211)

### Page 177

IN. MLE~ FLI T R

a. Exp Ianation.

(1) Each part of the rotorcraft must be free from excessive vibration under each appropriate speed and power condition (rule statement).

(2) This flight requirement may be both a qualitative and quantitative flight evaluation. Section 29.571(a) contains the flight load survey requirement that results in accumulation of vibration quantitative data. Section 29.629 generally requires quantitative data to show freedom from flutter for each part of the rotorcraft including control or stabilizing surfaces and rotors. See Paragraph No.’s 230 and 252 of this document for these two rules.

(3) Review Case No. 70 (reference FAA Order 81 10.6) contains a policy statement concerning compliance with this rule. This policy statement is condensed here for convenience:

“The rotorcraft must be capable of attaining a 30° bank angle (turn), at V~~, with maximum continuous power (maximum continuous torque) without encountering excessive roughness/vibration. The FA4/AUTHORITY requires the maneuver demonstration to provide the pilot with some maneuver capability at V~~, and further to provide the pilot some margin away from roughness when operating in turbulence.”

(This maneuver may result in a descent or a climb.)

(4) Section 29.1505 pertains to V~~ determination. Section 29.1509 pertains to rotor speed limits determination. See Paragraphs 720 and 721 of this document.

b.Jmxd!ms

(1) During the company flight test program, the rotorcraft is flown to the appropriate rotor and airspeed limits at several weights to prove that the rotorcraft is free from excessive vibration under appropriate speed, power, and weight conditions.

The flight loads survey quantitative data (reference ~ 29.571) and the applicant’s qualitative and quantitative flight test data must also prove compliance with the requirement prior to issuing an authorization for official FANAUTHORITY flight tests.

(2) The flight load survey data obtained under $ 29.571(a) will contain measured data concerning proof of freedom from flutter and excessive vibration.

Pertinent critical flight conditions will be reinvestigated during FAA/AUTHORITY flight tests. The specific condition or conditions necessary to demonstrate compliance with S 29.251 varies with the rotorcraft design, and with the minimum and maximum rotor 212 Par 110

### Page 178

speeds, VNE and VD speeds, and weight and CG position. An illustration of the speed and RPM demonstration is shown in Figure 110-1. Also see subparagraph b(4).

(3) The airspeed and rotor speed limits investigated and established under $~ 29.33,29.1503,29.1505, and 29.1509 are also investigated and made a matter of record in the flight loads survey data. During the official FAA/AUTHORITY/TIA flight tests, critical parts of the rotorcraft may have limited instrumentation to reinvestigate and confirm that the critical conditions investigated during the flight load survey are satisfactory and do not result in excessive vibration. Use of instrumentation is optional if the flight loads data, reference Paragraph No. 230 of this document, are conclusive.

(4) FAA policy for certification (Review Case No. 70) requires a “rotor roughness” flight demonstration of a 30° bank angle left and right, at maximum continuous power (MCP) (maximum continuous torque which may be in excess of the maximum continuous temperature limit), at VNE. To provide the pilot with some margin from roughness, the FAA requires maneuver demonstrations of 30 °banked turns at VN~ without encountering excessive roughness. The maneuver should be conducted with the rotor speed at the minimum RPM and maximum RPM limits. During the flight load survey, this condition should be investigated and data recorded to assure hazardous loads are not encountered for this “unusual” condition. As indicated, the flight condition will be reinvestigated during the FNVAUTHORITY flight tests. See Paragraph b(2) for illustration of this speed and RPM demonstration.

Par 110 213

### Page 179

AC 29-2B 7130197 0 L AI05?40Max RPM Power-Off ~ 4 ❑ IMax RPM Power-Off t t Max RPM Power-On ❑7 ROTOR RPM t Min RPM Power-On ❑“ 1,

2.

3.

4.

5.

6.

7.

8.

I Min RPM Power-Off t i 95V0 Min RPM Power-Off ~3 0° Bank, Autorotation ~ofi~rF~~rd Flight, 30° Bank, Power-On 02 ~6 A5 01 I I v NE 1.11 VNE

AIRSPEED Autorotation at 1.1I VNE(AR)at minimum Placard rotor sPeed” Autorotation at 1.I I NNE(AR)at maximum Placard rotor sPeed.

Autorotation at NN~(AR)at power-off minimum design limit rotor sPeed.

Autorotation at NNE~AR)at power-off maximum design limit rotor sPeed.

Forward flight 1.1 I VNEat minimum Power-on rotor sPeed.

Forward flight 1.1 I VNEat maximum Power-on rotor sPeed.

Right and left turn at VNEat maximum power-on rotor speed with 30° bank angle.

Right and left turn at VNEat minimum power-on rotor speed with 30° bank angle.

Note: VNE(AR)may be less than VNE” 214 FIGURE 110-1. DEMONSTRATION POINTS Par 110

### Page 180

7/30/97

**111.-119. RESERVED.**

Par Ill Intentionally Left Blank AC 29-2B 215 (thru 223)

### Page 181

AC 29-2B

120. KSl=O=.

224 7/30/97 Intentionally Left Blank Par 120

### Page 182

**SEC ION 7, STRFNGTT H REQUI REMFNTS - GENERAL**

121. ~ 29.301 LOADS,

a. danatio~.

(1) The rule is a general statement concerning limit and ultimate loads and the application of these loads to the rotorcraft.

(2) Ultimate loads are limit loads multiplied by the prescribed factors of safety.

(3) The specified loads must be distributed appropriately or conservatively and significant changes in distribution of the loads, as a result of deflection, must be taken into account.

b. Procedures. The design criteria report and/or design loads report must contain data that comply with the rule.

a. planatlon.

(1) Unless otherwise provided by FAR Part 29, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will assure that the design strength of the rotorcraft is greater than the design loads contained in FAR Part 29.

(2) Other rules, $$ 29.561(b)(3) and 29.787(c), specify use of defined ultimate inertia forces for protection of occupants.

b. Procedures.

(1) The design criteria report and/or design loads report must contain data that include the appropriate factor of safety.

(2) The factor of safety multiplies the limit external and inertia loads. The rule does allow the application of this factor to the resulting “limit internal” stresses if it is more conservative.

Par 121 225

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123. ~ 29.305 ST RENGTH AND DEFORMATION.

a. Explanation.

(1) This general rule defines, in relative terms, allowable deformation for limit and ultimate loads.

(2) If static tests are used to show compliance with this rule, the structure must support ultimate loads for 3 seconds without failure. Alternatively, dynamic tests simulating actual load applications may be used.

(3) Section 29.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by structural analysis. See Paragraph 124 of this AC.

b. Procedures. Any test results, static or dynamic, must satisfy the limitations or acceptance criteria contained in the rule.

(1) Any test proposals submitted for approval that are used to demonstrate compliance with sections of FAR Part 29 must contain the criteria stated in the rule.

(2) Any test results reports must contain data and information showing the test results comply with the standard.

(3) When dynamic tests are not used to substantiate the ultimate strength of structure subject to significant dynamic response under load, the analytical substantiation should consider flexibility effects and rate of load application (tail boom strength under landing loads is an example of a strength which needs dynamic amplification effects considered).

124. S 29 .307 (Amend ment 29-4) PROOF OF STRUCTURE.

a. Exdanation.

(1) The rule requires compliance for each critical loading condition. Certain tests must be conducted as specified. Additional tests for new or unusual design features may be required as noted in ~ 29.307(b)(6).

(2) “Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable.”

(3) Fatigue substantiation requirements are explained further in Paragraph 230 of this advisory circular.

b. Bwxh+ma F’ar 123

### Page 184

(1) The design criteria and/or design loads report should contain typical or representative loading conditions from which the critical loading conditions will be selected for analytical substantiation in structural (static and fatigue) reports and dynamics (vibration and stability) reports and fatigue, static, dynamic, or operational test reports.

(2) Whenever tests are used or required, a test proposal or plan must be approved prior to the tests. The test article must have received conformity inspections and must have been accepted by the FANAUTHORITY for the test. Test fixtures and instrumentation must also be acceptable to the FAA/AUTHORITY (using DERs as appropriate) prior to the start of the test. The quality control office of the applicant or other qualified personnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FANAUTHORITY or DER performing this task, The test proposal may be used to define and to authorize the means to accomplish inspection of the test fixtures and instrumentation. Unnecessary drawings, such as test fixture details, or layering of approvals is not intended or envisaged by this policy.

Drawings, sketches, or photographs have been used by the FAA/AUTHORITY to control and to assure correct location, direction, and magnitude of loads and other critical test parameters.

(3) Structural analysis has been accepted for rotorcraft in place of static tests.

Generally the rotorcraft airframe should have frequency placements remote to predominate rotor excitation sources, including rotor harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this vibration. During the flight load measurement program conducted under $29.571, critical Ioadecl areas or critical joints may be instrumented with strain gages or other stress strain measuring devices. This actual flight data may be compared to the analytical data to verify accuracy.

(4) Subparagraph (b) of the rule specifies certain tests. Test proposals must be approved prior to conducting official FWAUTHORITY tests. Other paragraphs in this advisory circular pertain to those tests.

124A. 629.307 (Amendment 79-30) PROO F OF STRU CTUK,

a. Ext)lanation, Amendment 29-30 adds the requirement to account for the environment to which the structure will be exposed in operation. This change is intended to codify recent FANAUTHORITY and industry practices for the consideration of environmental effects in showing “proof of structure.”

b. Procedures. All of the policy material pertaining to this section remains in effect with the following additions:

par 124 227

### Page 185

(1) For either tests or an analysis, environmental effects are now explicitly required. Consideration of loss of strength and stiffness of metals with elevated temperatures and loss of strength and stiffness of composite materials from exposure to heat, moisture, or other operational environments is now required and should be documented in analyses and test reports.

(2) MIL-HDBK-5D, AC 20-I07B, or MIL-HDBK-17B (or later versions) are acceptable sources of data and procedures to show compliance with environmental effects of metallic and composite materials, respectively.

125. ~ 29.309 DESIGN LIMITATIONS.

a. Exrdanation.

(1) The rule requires an orderly selection and presentation of the basic structural design limitations of the rotorcraft. The applicant must establish these structural limitations to facilitate design of the rotorcraft.

(2) Refer to the rule for the specific requirements.

b. Procedures.

(1) The design criteria and/or design load report should contain the design limits specified.

(2) These items are structural design limits. Other requirements may result in narrowing the ranges of type design limits or in reducing limits. It is not necessa~ to revise structural design criteria limits to agree with more conservative operational limits established during the certification program. The operational limits may be subsequently expanded by additional flight tests to agree with design limits.

**126.-135. RESERVED.**

228 (thru 238) Par 124A

### Page 186

**7130197 SEC TION 8. FLIGHT LOADS AC 29-2B**

**136.529.321 GENFR AL - FLIGHT LOADS.**

a, Explanat ion.

(1) The rule specifies the way the loads will be applied to the rotorcraft. It requires load analysis from minimum to maximum design weight. Any practical distribution of disposable loads must be included in the analysis.

(2) Paragraph (a) of the rule states: “The flight load factor must be assumed to act normal to the longitudinal axis of the rotorcraft and to be equal in magnitude and opposite in direction to the rotorcraft inertia load factor at the center of gravity.”

b. Procedures.

(1) Derivation of the flight loads is required by and specified in ~ 29.337 through $29.351. This rule requires flight load determination from minimum to maximum weight and for disposable loads.

(2) The application of the design loads derived from the flight load factor will be as specified. The flight loads analysis data must comply with the rule.

137. ~ 29,337 (Amendment 29-30) LIMIT MANF,U VFRING LOAD FACTOR.

a. Explanat ion. The rotorcraft must be designed and substantiated to load factors as specified to provide a minimum level of structural integrity of the rotorcraft airframe and rotors.

(1) A range of design positive load factors from +3.5 to +2.0 maybe used,

(2) A range of design negative load factors from -1.0 to -0.5 maybe used.

(3) Load factors inside the range of +3.5 to -1.0 maybe used provided the probability of exceeding the design load factors is shown by analysis and flight tests to be extremely remote, and the selected load factors are appropriate to each weight condition between design maximum and minimum weights.

(4) Load factors exceeding these “minimums” may be used.

b. FIQG@UM.

Par 136 239

### Page 187

(1) The applicant may elect to substantiate the rotorcraft for a design maneuvering load factor less than +3.5 and more than -1.0. Whenever this option is used, an analytical study and flight demonstration are required.

(i) The maximum positive design load factor is +3.5 generally at a weight below maximum gross weight. The maximu m thrust capability of the main rotor combined with incremental lift of wings or sponsons, if installed, results in a maximum design positive load factor. An example of a load factor - gross weight curve is shown in Figure 137-1. Note the minimum positive design load factor is +2.0 even though the required analysis and flight demonstration may prove the rotorcraft is not capable of achieving this load factor. This curve also illustrates compliance with ~ 29.337(b)(2) since the design load factor varies with gross weight.

(ii) The largest negative design load factor is -1 .0; however, several current rotorcraft designs are not capable of achieving a negative load factor.

Therefore, -0.5 has been an acceptable structural design negative load factor for certain rotorcraft designs.

(2) Whenever the applicant analytically substantiates the lower load factors allowed by ~ 29.337(b), the applicant must conduct the flight demonstration required by ~ 29.337 (b)(l ). The flight test personnel must determine that the demonstration is conducted in a manner to show that the probability of exceeding the selected design load factors, (those factors less than +3.5 and more than -1 .0) is extremely remote.

(See Order 8110.4, Paragraph 166c(2)(c)).

(3) A numerical value has not been assigned to “extremely remote” in this standard.

240 Par 137

### Page 188

+3.5 Represents curve of Design Load Factor NZ o ——— ——— — -1.0 . I Minimum 5 f Gross Weight Lbs. Maximum

LOAD FACTOR GROSS WEIGHT CURVE FIGURE 137-1 Par 137 241

### Page 189

AC 29-2B 7130/97

138. ~ 29.339 RESU LTANT LIMIT MANEUVERING LOADS.

a. Exdanat ion. The rule specifies or defines the application of rotor and lift surface loads to the rotorcraft.

(1) The design maneuvering load factors required by ~ 29.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift.

(2) The rules $j$ 29.321,29.337,29.341, and 29.351 all complement one another and result in the derivation of design flight loads that will be imposed to assure structural integrity of the rotorcraft.

(3) The following assumptions and conditions are specified in the rule.

(i) The rule requires application of appropriate loads at each rotor hub and auxiliary lifting surface.

(ii) Power-on and power-off flight with maximum design rotor tip speed ratio and specific conditions that must be considered.

(iii) Rotor tip speed ratio, defined in the rule, has been carried forward from the initial rotorcraft certification rules issued in 1946. The rotor tip speed ratio is a basic parameter used in calculating rotor aerodynamic forces.

b. Procedures.

(1) The rule specifies an acceptable assumption concerning application of the rotorcraft maneuvering loads.

(2) The rotor tip speed ratio is a parameter found in textbooks and other books such as NACA Report No. 716. The equation in the rule contains angle, “a.” Report No. 716 also defines angle, “a,” as the angle of attack of the rotor disk. This definition is more easily understood than the definition contained in the rule.

(3) The rotorcraft design loads are derived as prescribed by $j~ 29.321,29.337,

29.341, and 29.351. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule.

139. ~29.341 GU ST LOADS.

a. Expla nation.

(1) The rotorcraft must be substantiated for the loads derived from 30 feet per second vertical and horizontal gusts from hovering to 1.11 V~~; i.e., (V~).

242 par 138

### Page 190

(2) Gust loads for any vertical stabilizing surface should be derived for lateral or sideward gusts, as well as the head-on horizontal gusts. See Paragraph No. 159, ~ 29.413(a)(2) of this advisory circular.

(3) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward, as well as for head-on gusts. See Paragraph 159.

b. Procedures.

(1) Either sharp-edged (instantaneous) gusts or sharp-edged gusts modified by an alleviation (attenuation) factor may be used for calculating aerodynamic loads for the rotorcraft and any installed stabilizing surfaces. The following conditions may be used:

(i) Vertical gusts may be considered normal to the flight path of the rotorcraft except during hover or low speed flight (20 knots or less) when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.

(ii) For a vertical stabilizing surface, the horizontal gusts are normal to the flight path of the rotorcraft except during hover or low speed flight when the gusts may be assumed normal to the longitudinal axis of the rotorcraft.

(iii) A primary effect of encountering the gust is to change the lift of the rotors and rotorcraft surfaces. Of primary concern is the gust load or lift created by the main rotor or rotors. The lift increment of the horizontal stabilizing surface and fuselage are generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.

(iv) The rotorcraft shall be assumed in stabilized level flight prior to meeting the gust.

(v) The gust velocity may be assumed uniform across the rotorcraft.

(vi) Gust loads on the stabilizing surfaces are required as stated in Paragraph 159 of this advisory circular.

(2) The rotorcraft design maneuvering load factors may generally exceed the design gust load factors calculated in compliance with this rule. This maybe attributed to the small incremental change in lift due to the 30 FPS gust. Nonetheless, design gust loads for the rotorcraft shall be calculated as specified in the rule to assure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load factor.

(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHWNASA - Ames Specialist’s Meeting on Rotorcraft Par 139 243

### Page 191

AC 29-2B 7130/97 Dynamics, February 13-15, 1974. Thepaper, entitled, ’’Helicopter Gust Response Characteristics Including Unsteady Aerodynamics Stall Effects,” was written by P.J.

Arcidiacono, R.R. Berquist, and W.T. Alexander, Jr. References listed in the paper may be helpful also.

**140.629.351 YAWING CO NDITIONS.**

a, Exp lanatio~. The rule requires proof of a rotorcraft “structural” yaw or sideslip design envelope. This sideslip envelope must cover minimum forward speed or hover to V~ or V~~ , whichever is less. The rotorcraft must be structurally safe for the thrust capability of the directional control system.

(1) The rotorcraft structure must be designed to withstand the loads for the specified yaw conditions. The standard does not require a structural flight demonstration, It is a structural design standard.

(2) Maximum displacement of the directional control, except as limited by pilot effort (130 pounds; S 29.397(a)), is required for the conditions cited in the rule. A control system rate limiter or a yaw damper may be used. The total displacement is therefore a function of time as well as the maximum effort applied (130 pounds).

(i) At low airspeeds, 90° yaw (sideward flight) should be the design limit.

(ii) At high airspeeds, stabilized yaw angle (stabilized sideslip) must be substantiated as stated in the rule.

(iii) At high airspeeds, the maximum tail rotor thrust will be combined with the vertical (directional) stabilizer surface load, if a stabilizer is used, as specified by ~ 29.351(b)(l).

(iv) At high airspeeds, while the rotorcraft is in the sideslip condition, the directional control is then returned to the neutral position, attendant with the flight condition. The tail rotor thrust will be added to the restoring force of the vertical stabilizer.

(v) Both right and left yaw conditions should be proven.

(3) The tail rotor attachment structure must comply with $29.403.

(4) The vertical stabilizing surface must also comply with ~ 29.413.

b. J3QGQES.

(1) Many of the current single main rotor rotorcraft designs have vertical (directional) stabilizing surfaces. These surfaces may be solely vertical stabilizing fins 244 Par 139

### Page 192

as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Model FHI 100. The Hiller FHI 100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.

(i) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.

(ii) For vertical stabilizers that house the tail rotor output gearbox, such as the Hiller Model FHI 100, the tail surface air loads will add to or subtract from the tail rotor thrust according to the flight condition under consideration.

NOTE: For one example: At stabilized yaw to the right (left pedal depressed to limit) (~ 29.351(b)(2)), the tail rotor thrust moment should equal the restoring moment of the tail boom, vertical stabilizer and main rotor torque. As stated by ~ 29.351 (b)(3), the tail rotor thrust moment then is added to the vertical stabilizer restoring moment. The addition of tail rotor thrust ($ 29.351(b)(3)) and vertical stabilizer load is generally one of the critical design conditions for the fuselage/tail boom.

(iii) For vertical stabilizers or fins that have an offset incidence angle with respect to the rotorcraft axis, the vertical fin moment is added, or subtracted as applicable, to the tail rotor thrust moment. The condition stated in $ 29.351(b)(1) may result in adding the fin load to the tail rotor thrust.

(iv) Low airspeed maneuvers, such as sideward, rearward, and hover turns over a spot, typically impose insignificant aerodynamic loads on the fuselage and/or tail boom. The aerodynamic loads at VH or VNE, whichever is required, are generally the significant aerodynamic design loads.

(v) A rational assessment of the various yaw conditions maybe used to reduce the load deviation and analysis to the critical rotorcraft design conditions.

(vi) The rotorcraft structure shall be analyzed or tested for loads derived from the critical design conditions.

(vii) A simple structural design envelope maybe derived from these design data. If the right or left yaw limits are not very different, common, conservative design limits may be used. A sample yaw/forward speed diagram, as derived from design analysis of the characteristics of a hypothetical rotorcraft, is presented in Figure 140-1. A table of values would also suffice. This figure reflects characteristics which include a 90° yaw when the directional control inputs are applied at low airspeeds (up to 30 knots presumably the maximum sideward flight speed of which this aircraft is capable) and 10° yaw when they are applied at VH, with a straight line variation from 30 knot forward speed to VH.

Par 140 245

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/ 90° 90” YAW DEGREES 10° MAX.

().

AIRSPEED 30 KNOT LESSER OF SIDEWARD V~ OR V~~ FLIGHT FIGURE 140-1 SAMPLE YAW/FORWARD SPEED DIAGRAM 246 Par 140

### Page 194

(viii) During flight test evaluations, yaw angles have been measured using a yaw angle probe (swiveling vane type) on a nose boom. Both q visual readout for the pilot and a record, such as an oscillograph trace, have been used. This test may be conducted in the flight test program or in the flight load survey program. This record should confirm the yaw angle used in design as conservative with respect to operational and actual flight characteristics. This test is not a requirement however.

140A. 29.351 (Amendment 79-30) YAWING CO NDITIONS.

lanatio~ Amendment 29-30 adds maximum sideslip angles to the existing structural design purposes. The standard should apply to power-on conditions; not power-off, since VH is a part of the standard. For airspeeds up to 0.6 V~~, sideslip angles larger than 90° (or sideward flight) need not be considered. For airspeeds at VNE or VH (whichever is less), sideslip angles larger than 15° need not be considered.

b. Procedu res.

(1) All of the policy material pertaining to this section remains in effect with the addition of the maximum sideslip limits of 90° and 15° specified above. The rotorcraft does not need to be capable of attaining these conditions. A revised yaw/fomvard speed diagram is presented in Figure 140-2.

(2) FAR ~ 29.351(b)(1) incorrectly references $ 29.395(a) for maximum pilot forces. The correct reference should be ~ 29.397(a).

Par 140 247

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90° v

YAW DEGREES (LONGITUDINAL Axis) “ 15° MAX.

0°

0.6 V~~ AIRSPEED 7

**LESSER OF**

VH OR VNE FIGURE 140-2. SAMPLE YAW/FORWARD SPEED DIAGRAM 248 Par 140

### Page 196

140B. 29,35 1 (Amendment 29-40) YAWING CONDITIONS.

a. Explanation. Amendment 29-31 deleted FAR ~ 29.403 and ~ 29.413 since the references and requirements are adequately addressed in ~~ 29.337, 29.339, and

29.341. Paragraphs 157 and 159 of this AC are retained as guidance information.

b. Procedures. All of the policy material pertaining to this section remains in effect except for references to $529.403 and 29.413 in Paragraphs 140a(3) and (4) of this AC.

141.$29,361 (Amendme nt 29-26) ENGINE TORQUE.

a. Exp Ianation.

(1) The rotorcraft should be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. The rule recognized that reciprocating (piston) engines generate higher torque oscillations than turbine engines.

(i) A factor of 1.25 applies to maximum continuous power for turbine engines. Section 29.923 refers to torque output and ~ 29.927(b) refers to other torque output conditions for use in an “endurance test.”

(ii) Torque factors are also specified for reciprocating engines having two or more cylinders in ~ 29.361(a)(2) or ~ 29.361 (b) of Amendment 29-26. The appropriate torque factor applies to takeoff power torque as well as maximum continuous power and other power conditions.

(2) Amendment 29-26 introduced additional turbine engine installation considerations for the following:

(i) Engine torque loads associated with emergency operation of governor-controlled turboshaft engines.

(ii) Torque reaction loads from sudden turbine engine stoppage which is applied to the engine and the engine suspension and restraint system.

(3) Paragraph 206 of this document concerns $ 29.549(c) and (e) that contains design standards for engine mounts and adjacent structure for flight and landing and also flight with 2 %-minute OEI power rating. Amendment 29-26 added OEI power to the standard.

(4) Section 29.547 (e)(l)(ii) concerns the application of limit engine torque to design of the main rotor structure.

Par 140B 249

### Page 197

AC 29-26 7130197

b. Procedures.

(1) The engine torque associated with the maximum continuous power condition should be multiplied by the appropriate torque factor to obtain the engine torque value used for structural substantiation purposes of the rotorcraft.

(2) The torque values associated with the minimum power-on RPM limit should be used. Maximum power-on speed limit will result in a lower torque value when calculating torque from design horsepower values. However, due to piston engine power output characteristics, an engine may produce a higher torque at higher engine speeds contrary to the previous statement. The torque factor should account for this characteristic.

(3) For turbine engines limit torque values are determined for the four cases cited. Two cases are related to “endurance” test standards.

(4) For sudden stoppage of turbine engines the engine manufacturer can reasonably provide engine rotating inertia and deceleration time expected in the event of sudden engine stoppage which generates these critical loads in the engine mounting and restraint system. These manufacturer’s data should be acceptable for use in complying with this part of the design standard.

**142.-151. RESER VED.**

250 (thru 260) Par 141

### Page 198

**SEC TION 9, CON TROL SURFAC E AND SYSTEM LOADS**

152. ~ 29.391 GENER AL.

a. Exdanat ion. This general rule concerns requirements for design loads of tail rotors, control or stabilizing surfaces, and their control system.

b. procedures. The design criteria and/or the design loads report must contain the loads dictated by the referenced rules. See Paragraphs 153, 154, 155, 156, 157, 158, and 159.

152A. $29,391 (Amendment 29-30) GEN ERAL.

a. ~XP anatI ion. Amendment 29-30 adds an explicit reference to ~ 29.427, U svm metrical Load s (AC Paragraph 162), to clarify that substantiation for u~symmetrical loads is a general control sutiace requirement. A reference to ~ 29.399, Dual Contro SvsteI m (AC Paragraph 155), is also added for clarification. In addition, $j~ 29.401,29.403,29.413 were removed by this amendment since these references and requirements were adequately addressed in other standards.

b. Procedures. The referenced AC paragraphs become 153, 153A, I 54, 155, 158, and 162.

153. ~ 29.395 CO NTRO L SYSTEM.

a. Exrdanation. Control system design loads and the application of these loads are contained in this rule.

(1) Paragraph (a) of the rule specifies the way or means of reacting the design loads specified in $j~ 29.397 and 29.399 (for dual control systems). The design loads must be imposed on any locks and stops and irreversible mechanisms in the control system. Both rotor blade horns and control surface horns must react without failure, the specified loads while the controls are in critical positions.

(2) Paragraph (b) of the rule specifies application of limit pilot forces or of the maximum loads that can be obtained in normal operation, including any single power boost system failure, whichever is greater. However, minimum limit pilot force 0.60 of the loads specified in $j~ 29.397 and 29.399, may be used, as specified, in parts of the primary control system that are not stiff enough to react to the loads specified in the first part of Paragraph (b) of the rule. Note the objective for a rugged control system.

(3) Control system design feature and test requirements are found in ~!j 29.671 through 29.695. Bearing factors and fitting factors are specified in ~~ 29.623 and

29.625, respectively.

Par 152 261

### Page 199

b. Procedures.

(1) The design criteria and/or a design loads report that includes the primary control system design loads should be submitted for FAIVAUTHORITY approval.

(2) The rotorcraft control system maybe tested to ultimate design loads or may be analyzed for the ultimate design loads. See Paragraph 124 of this document.

(i) It is advisable that the applicant prepare a proposal describing the procedures and techniques to be used in the static testing of the control system which reflects compliance with the condition specified. It is further advisable that the FAWAUTHORITY concur that the tests proposed achieve that objective. Omission of these steps may result in the need for retesting. The test results should be documented.

(ii) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from ~~ 29.685(e), 29.623, and 29.625 must be used as specified. A structural analysis report should be used to document compliance with ~ 29.685(d)(1) and (4), and $j 29.685(f).

(3) If a part of the control system is not stiff or rigid enough to react the design loads specified in $29.397, that pant of the system maybe substantiated for lower loads as prescribed.

(i) The limit design loads are those loads specified in S 29.397;

(ii) The limit design loads are the maximum that can be obtained in normal operation, including any single power boost system failure, except for objectives stated for a rugged system; and

(iii) In lieu of a rational analysis, the limit design loads may be 0.60 of the loads specified in S 29.397.

(iv) For example, if a control surface servo tab or a small elevator is a part of the rotorcraft design, the control system for this part must be stiff enough to react the control surface loads without failure and to provide enough surface deflection to control the rotorcraft. These limit loads maybe 60 pounds fore and aft and 40 pounds laterally on the cyclic control stick in lieu of a rational analysis and may be the maximum loads that can be obtained in normal operation.

(v) If a hydraulic power actuation or boost system is part of the rotorcraft design, the design limit load for the affected parts of the control system will be the maximum output force of the boost at normal operating pressure added to the limit design loads resulting from the loads specified in ~ 29.397. If a single failure in the power portion of the hydraulic system results in actuator forces that exceed the 262 Par 153

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maximum output force at normal operating pressure, the highest output loads must be used as noted in Subparagraph (3)(ii). This hydraulic system failure standard is specified in ~ 29.695(a)(1) as well,

(4) Controls proof and operation test is required by ~fj 29.307(b), 29.681, and

29.683. This test is conducted using the design limit loads approved under S 29.395(b).

See Paragraphs 282 and 283 of this document.

153A. 29.395 (Amendment 29-30) CONTROL SYSTF~.

a. E&planation. Amendment 29-30 clarifies that the loads in 5 29.395(b) apply to power “control” systems not just power “boost” systems; and the limit pilot forces prescribed in ~ 29.397 are required to be applied in conjunction with the forces from normally energized power devices. The amendment may increase required loads for systems if operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. If so, the system is required to withstand 100 percent of limit pilot forces specified in ~ 29.397, rather than 60 percent of the limit pilot forces as specified previously.

b. procedures. The procedures of Paragraph 153 continue to apply except that the increased loads in new Paragraph ~ 29.355(b)(4) of 100 percent of limit pilot forces are specified for systems where operational loads may be exceeded by jamming, ground gusts, control inertia, or friction.

154.$29.397 (Amendme nt 29-12) LiMIT PILOT FORCES AND TORQU ES.

a. Exdanation. Design forces are contained in the rule.

(’l) Primary controls, pilot and copilot, must be designed for the limit pilot forces specified in Paragraph (a) of the rule.

(2) For other operating controls, such as flap, tab, stabilizer, rotor brake, and landing gear, design limit forces are specified in Paragraph (b).

b. procedures.

(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis of the control systems submitted in compliance with other rules.

(2) Operation tests of the control systems noted in other rules require application of these forces also.

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AC 29-2B 7130197

**155.629.399 DUAL CONT ROL SYSTEM..**

a. Design limit loads are specified for dual control systems. Pilot effort forces applied in opposition and in the same direction are required for dual control systems.

b. PIQwams

(1) Design loads specified in the rule must be used in required structural tests and in any structural strength analysis submitted for compliance with the other rules.

(2) Operation tests of the control systems, noted in other rules, require application of these forces also.

156. ~ 29.401 (Amendment 29 -4) AU XILIARY ROTOR ASSE MBLIES.

a. Explanat ion.

(1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by $29.923 and structural strength substantiation is required. Section 29.401 (b) specifically refers to structural strength substantiation for centrifugal loads resulting from maximum design rotor RPM. Due to the pitch feathering requirements, auxiliary rotors typically have detachable blades.

(2) The rotor blade structure must have sufficient strength to withstand not only aerodynamic loads generated on the blade surface, but also inertial loads arising from centrifugal, coriolis, gyroscopic, and vibratory effects produced by this blade movement.

Sufficient stiffness and rigidity must be designed into the blades to prevent excessive deformation and to assure that the blades will maintain the desired aerodynamic characteristics. As a design objective, the structural strength requirements should be met with the minimum material. Excess blade weight imposes extra centrifugal loads that may increase the operating stress levels. Blade weight and strength should be optimized. Even though a structural strength analysis for the blade design loads is required, a flight load survey and fatigue analysis are also required by 529.571.

(3) Section 29.1509 defines the design rotor speed as that providing a 5 percent margin beyond the rotor operating speed limits.

b. Procedu res.

(1) The endurance tests prescribed by !j~ 29.923 and 29.927 require achieving cettain speeds, power, and control displacement for the auxiliary (tail) rotor as well as the main rotor. The parts must be serviceable at the conclusion of the tests.

264 Par 155

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7130197 AC 29-2B

(2) Structural substantiation of the auxiliary (tail) rotor is required to assure integrity for the minimum and maximum design rotor speeds and the maximum design rotor thrust in the positive and negative direction. Thrust capability of the rotor should offset the main rotor torque at maximum power as required by !$29.927(b).

(i) The maximum and minimum operating rotor speed, power-off, is 95 percent of the maximum design speed and is 105 percent of the minimum design speed, respectively.

(ii) The rotor operating speed limits shown during the official FAWAUTHORITY flight tests must include the noted 5 percent margin with respect to the design speeds.

(iii) The auxiliary rotor generally has a positive and negative pitch limit that assures adequate directional control throughout the operating range of the rotorcraft. The power-off rotor speed limits are generally broader than the power-on rotor speed limits because of the required autorotational rotor speed characteristics.

Thus, the auxiliary rotor design conditions concern the maximum and minimum design rotor speeds in conjunction with the maximum positive or negative pitch thrust as appropriate. Thrust capability and precone angle of the rotor, if any, will significantly influence the rotor design loads. The variations in rotor design features and an example of substantiation would be too lengthy to include here. However, ANC-9, ‘tAircraft Propeller Handbook,” contains principles that may be applied to tail rotor designs. Tail rotors may be considered a special propeller design.

(iv) Bearings are generally used in the tail rotor installation to allow flapping and feathering motion of the blades. The bearings manufacturer’s ratings of these bearings must not be exceeded. Bearings generally used in main and tail rotors are classified as ABEC Class 3, 5, or 7. Class 7 is the highest quality presently available. Satisfactory completion of the endurance tests of $$29.923 and 29.927 is a means of proving that use of a particular bearing is satisfactory.

(v) The analysis must include appropriate special factors, casting factors, bearing factors, and fitting factors prescribed by ~~ 29.619, 29.621, 29.623, and

29.625, respectively. The fitting factor of 1.15 must be applied in the analysis of the tail rotor installation.

156A. 629.401 (Amendment 29-31) AUXILIARY ROTOR ASSEMBLIES,

a. Explanat ion. Amendment 29-31 removed this section since the requirements are adequately addressed in ~~ 29.337, 29.339, and 29.341.

b. Procedures. The policy material pertaining to this section is retained as supplemental information.

Par 156 265

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157. ~,

a. tio~.

(1) The auxiliary rotor attachment structure(s), which is considered to include gear boxes, must be designed to withstand design limit loads that occur in flight and on landing, These design loads that generally consist of the following must be established for the particular flight and landing condition under consideration.

(i) Inertia loads generated by linear and angular accelerations of the auxiliary rotors and their gear boxes, combined with

(ii) Thrust and torque loads developed by the auxiliary rotors.

The linear and angular acceleration loads imposed by the weight of the tail rotor and gearbox are generally derived from airframe loads data. Thrust and torque output of the tail rotor are derived during external aerodynamic and landing loads development for pertinent flight and landing conditions.

(2) General rules related to proof of structure loads and factor of safety are 5$29.307,29.301,29.303, and 29.305.

b. B.wdLEs

(1) The angular and linear acceleration loads combined with appropriate tail rotor thrust and torque for the critical conditions shall be imposed on the tail rotor gearbox mount lugs, the airframe mounting structure, and the attaching hardware.

(2) The yaw and maximum power climb conditions are generally critical.

Landing and maneuvering conditions with and without power may also impose high inertia and rotor thrust and torque loads on the attachment structure.

(3) The derivation of the loads and conditions are too extensive to include here.

Additional information can be found in the U.S. Army Material Command Report AMCP 706-201, “Engineering Design Handbook: Helicopter Engineering, Part One, Preliminary Design.”

157A. ILIARY ROTO R ATTACHMENT planat ion. Amendment 29-31 removed this section since the requirements are a~equately addressed in ~$j 29.337, 29.339, and 29.341.

b. procedures. The policy material pertaining to this section is retained as supplemental information.

266 Par 157

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