# AC 27-1A: Certification of Normal Category Rotorcraft

- Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it)
- Status: Historical
- Version in force: [AC 27-1B CHG 1-9](https://openairdocs.org/ac/27-1), approved June 23, 2023. This is a historical version
- Approved: July 30, 1997
- Last reviewed: January 23, 2001
- Family: AC 27-1
- Cancelled by: [AC 27-1B](https://openairdocs.org/ac/27-1B?v=DRSDOCID123898489720230413220327)
- 14 CFR parts: 27
- Pages: 690
- PDF: https://openairdocs.org/files/sha256/42/4b/424bdbe333f60ce7865a2467b89831737568bb6ba59aeca9e39c958e4ec5abed
- FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/6076DA5331CE8D8B862569DD005BC2B2.0001
- DRS record last changed: November 22, 2025
- Page: https://openairdocs.org/ac/27-1A

## Versions

- [AC 27-1B CHG 1-9](https://openairdocs.org/ac/27-1) (Current), approved Jun 23, 2023, 1,196 pages
- [AC 27-1B](https://openairdocs.org/ac/27-1B?v=DRSDOCID123898489720230413220327) (Historical), approved Jun 23, 2023, 1,196 pages
- [AC 27-1B CHG 1-8](https://openairdocs.org/ac/27-1B-CHG-1-8) (Historical), approved Jun 29, 2018, 1,140 pages
- [AC 27-1B CHG 1-2](https://openairdocs.org/ac/27-1B-CHG-1-2) (Historical), approved Apr 25, 2006, 932 pages
- [AC 27-1B](https://openairdocs.org/ac/27-1B?v=0CE1B2A7CEE21AAE8625703800684002) (Historical), approved Jun 24, 2005, 7 pages. CHGS to Various Paragraphs
- [AC 27-1B](https://openairdocs.org/ac/27-1B?v=771532252D914FF686256D42006468E0) (Historical), approved Apr 25, 2003, 3 pages
- [AC 27-1B](https://openairdocs.org/ac/27-1B?v=897391F05A323C6186256CEE005D87AF) (Historical), approved Feb 12, 2003, 2 pages
- [AC 27-1B](https://openairdocs.org/ac/27-1B?v=6F95F1C7B8B1E89C86256E4300597919) (Historical), approved Feb 12, 2003, 901 pages
- [AC 27-1A](https://openairdocs.org/ac/27-1A) (Historical), approved Jul 30, 1997, 690 pages
- [AC 27-1B](https://openairdocs.org/ac/27-1B?v=5232FE84F9AF806C8625707B007183F6) (Historical), approved , 4 pages. 27.351 - Yawing Conditions

## ADs that cite it (1)

- [AD 2018-02-01](https://openairdocs.org/ad/2018-02-01): Main Rotor Head · Enstrom Helicopter Corporation, The 280, 280C, 280F and 8 more. Effective Feb 21, 2018

## Text

Each page's text from the PDF; open the PDF at a page with #page=N. This is OCR from the scanned FAA advisory circular. OCR can misread letters and numbers.

### Page 1 (OCR)

Advisory of Transportation U.S. Department crau ar Federal Aviation Administration Subject: CERTIFICATION OF NORMAL Date: 7/30/97 AC No: 27-1A CATEGORY ROTORCRAFT Initiated by: ASW-110 Change:

1. PURPOSE:

a. This is a total revision of AC 27-1 dated 8/29/85, with changes 1, 2, 3, and 4 dated 9/16/87, 4/24/89, 9/12/91, and 8/18/95 respectively, incorporated. In addition, new material plus changes to existing paragraphs have been incorporated. This consolidated version is now renumbered as AC 27-1A and replaces AC 27-1 in its entirety. This revises existing material in 14 paragraphs and adds new material for 17 paragraphs.

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 27, which contains the Airworthiness Standards for Normal Category Rotorcraft. It includes methods of compliance in the areas of basic design, ground tests, and flight tests.

2. CANCELLATION. AC 27-1, Certification of Normal Category Rotorcraft, August 29, 1985, 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 32, 45, 58, 338, 532, 619, 621, 633, 641, 720, 726, 765, 775, and 777 are revised to incorporate technical guidance.

b. New paragraphs 31A, 58B, 63A, 218B, 218C, 230A, 423B, 447, 454B, 456, 459A, 460B, 563B, 568, 619B, and 724B are added to Chapter 2.

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

d. Paragraph 447, § 27.951, General, is renumbered to Paragraph 446. Paragraph 447 now addresses § 27.952, Fuel System Crash Resistance.

For sale by the U.S. Govemment Printing Office Superintendent of Documents, Mail Stop: SSOP, Washington, DC 20402-9328 ISBN 0-16-049210-6

### Page 2 (OCR)

e. The following appendix has been added:

Appendix 1 Rotorcraft One-Engine-Inoperative (OEI) Power Assurance

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

5. DEVIATIONS. As rotorcraft designs vary from conventional configurations, it may become necessary to deviate from the methods and procedures outlined in this AC. These procedures are only one acceptable means of compliance with Part 27. Any alternate means proposed by an applicant will be given due consideration. Applicants are encouraged to use their technical ingenuity and resourcefulness to develop more efficient and less costly methods of achieving the objectives of Part 27. Regulatory personnel and designees should respond to such efforts by the use of engineering judgment in fostering any such efforts as long as the letter and spirit of Part 27 and the Federal Aviation Act are respected. It is recommended that unusual or unique projects be coordinated a sufficient time in advance with the Rotorcraft Standards Staff, ASW-110, or with the appropriate airworthiness authority, to ensure timely and uniform consideration.

6. APPLICABILITY. This material is not to be construed as having any legally binding status and must be treated as advisory only. However, to ensure standardization in the certification process, these procedures should be considered during all rotorcraft type certification and supplemental type certification activities.

7. PARAGRAPHS KEYED TO FAR PART 27. Each paragraph has the applicable amendment to Part 27 shown in the title. All of the original guidance material has been retained as appropriate, even as changes are made to the regulations. This is accomplished through the use of "A," "B," etc., paragraphs 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 (OCR)

TABLE OF CONTENTS CHAPTER 1 - FAR 21 CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS Page No.

**1.-3. RESERVED**

4. FAR 21.16 Special Conditions. 1

**5.-7 RESERVED**

8. FAR 21.31 Type Design.. 6

**9.-11. RESERVED**

12. FAR 21.33 Inspection and Tests.. .. 10

**13.-15. RESERVED**

16. FAR 21.35 Flight Tests.. . 14

**17.-23. RESERVED**

24. FAR 21.39 Flight Test Instrument. .. 24 Calibration and Correction Report

25.-30. RESERVED CHAPTER 2 - FAR 27 AIRWORTHINESS STANDARDS: NORMAL CATEGORY

**ROTORCRAFT SECTION 1. GENERAL**

31. FAR 27.1 Applicability 31 31A. FAR 27.1 Applicability 31

32. FAR 27.2 Special Retroactive Requirements 31

**33.-41. RESERVED SECTION 2. FLIGHT - GENERAL**

42. FAR 27.21 Proof of Compliance. 43

43. FAR 27.25 Weight Limits. 44

44.  FAR 27.27 Center of Gravity Limits … 46

45. FAR 27.29 Empty Weight and Corresponding Center of Gravity. 48

46. FAR 27.31 Removable Ballast 49

47. FAR 27.33 Main Rotor Speed and Pitch Limits . 49

**48.-57. RESERVED SECTION 3. PERFORMANCE**

58. FAR 27.45 General ..61 58A. FAR 27.45 General ... 70 58B. FAR 27.45 General .. 70

59. FAR 27.51 Takeoff .77

### Page 4 (OCR)

SECTION 3. PERFORMANCE (continued) Page No.

**60.-62. RESERVED**

63. FAR 27.65 Climb: All Engines Operating. .. 82 63A. FAR 27.65 Climb: All Engines Operating.. .. 84

64. FAR 27.67 Climb: One Engine Inoperative.. 84

65. FAR 27.71 Glide Performance ............. ..85

66. FAR 27.73 Performance at Minimum Operating Speed. .. 88

67. FAR 27.75 Landing ..... .. 90

**68. RESERVED**

69. FAR 27.79 Limiting Height - Speed Envelope 95 69A. FAR 27.79 Limiting Height - Speed Envelope 103

**70.-79. RESERVED SECTION 4. FLIGHT CHARACTERISTICS**

80. FAR 27.141 General. . 114

81. FAR 27.143 Controllability and Maneuverability .. .. 115

82. FAR 27.151 Flight Controls 120

83. FAR 27.161 Trim Control .. 121 84 FAR 27.171 Stability: General 122

85. FAR 27.173 Static Longitudinal Stability.. . 122 86 FAR 27.175 Demonstration of Static Longitudinal Stability. 123

87. FAR 27.177 Static Directional Stability ... 128

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

96. FAR 27.231 General . 138

97. FAR 27.235 Taxiing Condition. 138

98. FAR 27.239 Spray Characteristics 139

99. FAR 27.241 Ground Resonance 140

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

110. FAR 27.251 Vibration. . 155

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

121. FAR 27.301 Loads. 168

122. FAR 27.303 Factor of Safety.. 168 IV

### Page 5 (OCR)

SECTION 7. STRENGTH REQUIREMENTS - GENERAL (continued) Page No.

123. FAR 27.305 Strength and Deformation 168

124. FAR 27.307 Proof of Structure. 169 124A. FAR 27.307 Proof of Structure. 170

125. FAR 27.309 Design Limitations. 170

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

136. FAR 27.321 General .. 182

137. FAR 27.337 Limit Maneuvering Load Factor. 182

138. FAR 27.339 Resultant Limit Maneuvering Loads 185

139. FAR 27.341 Gust Loads. 185

140. FAR 27.351 Yawing Conditions. 186

141. FAR 27.361 Engine Torque. 190

**142.-151. RESERVED SECTION 9. CONTROL SURFACE AND SYSTEM LOADS**

152. FAR 27.391 General . 202 152A. FAR 27.391 General. 202 FAR 27.395 Control System 202 153A. FAR 27.395 Control System. 204 FAR 27.397 Limit Pilot Forces and Torques 204 FAR 27.399 Dual Control System. 205 FAR 27.401 Auxiliary Rotor Assemblies 205 156A. FAR 27.401 Auxiliary Rotor Assemblies. 207 FAR 27.403 Auxiliary Rotor Attachment Structure. 207 157A. FAR 27.403 Auxiliary Rotor Attachment Structure 208

158. FAR 27.411 Ground Clearance: Tail Rotor Guard 208

159. FAR 27.413 Stabilizing and Control Surfaces 209 159A. FAR 27.413 Stabilizing and Control Surfaces.. ..211 FAR 27.427 Unsymmetrical Loads.. 211

**161.-169. RESERVED SECTION 10. GROUND LOADS**

170. FAR 27.471 General .. 225

171. FAR 27.473 Ground Loading Conditions and Assumptions. 225

172. FAR 27.475 Tires and Shock Absorbers 226

173. FAR 27.477 Landing Gear Arrangement 226

174. FAR 27.479 Level Landing Conditions. 226

### Page 6 (OCR)

SECTION 10. GROUND LOADS (continued) Page No.

175. FAR 27.481 Tail-Down Landing Conditions 227

176. FAR 27.483 One-Wheel Landing Conditions.. .. 228

177. FAR 27.485 Lateral Drift Landing Conditions. .. 228

178. FAR 27.493 Braked Roll Conditions....... .. 229

179. FAR 27.497 Ground Loading Conditions: 229 Landing Gear With Tail Wheels

180. FAR 27.501 Ground Loading Conditions 231 Landing Gear With Skids 180A. FAR 27.501 Ground Loading Conditions: 232 Landing Gear With Skids

181. FAR 27.505 Ski Landing Conditions... 233

**182.-192. RESERVED SECTION 11. WATER LOADS**

**193. RESERVED**

194. FAR 27.521 Float Landing Conditions.. 245

**195.-204. RESERVED SECTION 12. MAIN COMPONENT REQUIREMENTS**

205. FAR 27.547 Main Rotor Structure. 258

206. FAR 27.549 Fuselage, Landing Gear, and Rotor Pylon Structure 258

**207.-217. RESERVED SECTION 13. EMERGENCY LANDING CONDITIONS**

218. FAR 27.561 General .. 271 218AFAR 27.561 Emergency Landing Conditions - General.. .272 218B. FAR 27.561 Emergency Landing Conditions - General 272 218C. FAR 27.561 Emergency Landing Conditions - General.. .. 273

219. FAR 27.562 Emergency Landing Dynamic Conditions. 273

220. FAR 27.563 Structural Ditching Provisions.. .274 220A. FAR 27.563 Structural Ditching Provisions. . 274

**221.-229. RESERVED SECTION 14. FATIGUE EVALUATION**

230. FAR 27.571 Fatigue Evaluation of Flight Structure. .. 285 230A. FAR 27.571 Fatigue Evaluation of Flight Structure for Category A Certification 288

**231.-240. RESERVED VI**

### Page 7 (OCR)

SECTION 15. DESIGN AND CONSTRUCTION - GENERAL Page No.

241. FAR 27.601 Design. .. 300

242. FAR 27.603 Materials .. . 300

243. FAR 27.605 Fabrication Methods.. 301

244. FAR 27.607 Fasteners.. 302

245.  FAR 27.609 Protection of Structure … 302

246. FAR 27.610 Lightning Protection... 303

247. FAR 27.611 Inspection Provisions ..... .. 306

248. FAR 27.613 Material Strength Properties 306 and Design Values 248A. FAR 27.613 Material Strength Properties .308 and Design Values

249. FAR 27.619 Special Factors . .. 309

250. FAR 27.621 Casting Factors 309

251. FAR 27.623 Bearing Factors .. 319

252. FAR 27.625 Fitting Factors .. 320

253. FAR 27.629 Flutter.. 321

**254.-264. RESERVED SECTION 16. ROTORS**

265. FAR 27.653 Pressure Venting and Drainage of Rotor Blades . 334

266. FAR 27.659 Mass Balance. 334

267. FAR 27.661 Rotor Blade Clearance... .. 334

268. FAR 27.663 Ground Resonance Prevention Means.. 335 268A. FAR 27.663 Ground Resonance Prevention Means. .. 337

**269.-277. RESERVED SECTION 17. CONTROL SYSTEMS**

278. FAR 27.671 General .. 348

279. FAR 27.672 Stability Augmentation, Automatic,. .. 348 and Power-Operated Systems FAR 27.673 Primary Flight Control..... .349 FAR 27.674 Interconnected Controls. .. 349 FAR 27.675 Stops.. . 350

FAR 27.679 Control System Locks … 351

FAR 27.681 Limit Load Static Tests... 352 FAR 27.683 Operation Tests. 353 FAR 27.685 Control System Details... .. 353 286A. FAR 27.685 Control System Details... .354 FAR 27.687 Spring Devices.. 355

288. FAR 27.691 Autorotation Control Mechanism 355 VII

### Page 8 (OCR)

SECTION 17. CONTROL SYSTEMS (continued) Page No.

289. FAR 27.695 Power Boost and Power-Operated Control System ... 356

**290.-297. RESERVED SECTION 18. LANDING GEAR**

298. FAR 27.723 Shock Absorption Tests. .. 367

299. FAR 27.725 Limit Drop Test.. . 370

300. FAR 27.727 Reserve Energy Absorption Drop Test. .. 370 300A. FAR 27.727 Reserve Energy Absorption Drop Test.. ... 371

301. FAR 27.729 Retracting Mechanism ... 371

302. FAR 27.731 Wheels .. .. 374

303. FAR 27.733 Tires. 375

304. FAR 27.735 Brakes. 376

305. FAR 27.737 Skis 378

**306.-315. RESERVED SECTION 19. FLOATS AND HULLS**

316. FAR 27.751 Main Float Buoyancy 390

317. FAR 27.753 Main Float Design .. 390

318. FAR 27.755 Hulls. 391

319.-329. RESERVED SECTION 20. PERSONNEL AND CARGO ACCOMMODATIONS FAR 27.771 Pilot Compartment. ,404 FAR 27.773 Pilot Compartment View .405 FAR 27.775 Windshields and Windows. .410 332A. FAR 27.775 Windshields and Windows. .410 FAR 27.777 Cockpit Controls.. .410 FAR 27.779 Motion and Effect of Cockpit Controls ... ..411 FAR 27.783 Doors .414 335A. FAR 27.783 Doors .. .416

336. FAR 27.785 Seats, Berths, Safety Belts, and Harnesses .416 336A. FAR 27.785 Seats, Berths, Safety Belts, and Harnesses . 420

337. FAR 27.787 Cargo and Baggage Compartments.. . 422 337A. FAR 27.787 Cargo and Baggage Compartments. .424

338. FAR 27.801 Ditching. 424

**339. RESERVED**

340. FAR 27.807 Emergency Exits. 433 340A. FAR 27.807 Emergency Exits 434 VIlI

### Page 9 (OCR)

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

**341.-344. RESERVED**

345. FAR 27.831 Ventilation 439

346. FAR 27.833 Heaters .. .440

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

**357. RESERVED**

FAR 27.853 Compartment Interiors … 454

FAR 27.855 Cargo and Baggage Compartments .455

360. FAR 27.859 Heating Systems.. 456

361. FAR 27.861 Fire Protection of Structure, Controls, and Other Parts. 460 361A. FAR 27.861 Fire Protection of Structure, Controls, and Other Parts 462

362. FAR 27.863 Flammable Fluid Fire Protection 463

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

373. FAR 27.865 External Load Attaching Means. 475 373A. FAR 27.865 External Load Attaching Means 476

**374.-383. RESERVED SECTION 23. MISCELLANEOUS (DESIGN AND CONSTRUCTION)**

384. FAR 27.871 Leveling Marks.. .487

385. FAR 27.873 Ballast Provisions.. 487

**386.-396. RESERVED SECTION 24. POWERPLANT - GENERAL**

397. FAR 27.901 Installation.. .. 500 397A. FAR 27.901 Installation.. 503

398. FAR 27.903 Engines.. .. 504 398A. FAR 27.903 Engines... .506

399. FAR 27.907 Engine Vibration.... . 508

**400.-420. RESERVED IX**

### Page 10 (OCR)

SECTION 25, ROTOR DRIVE SYSTEM Page No.

421. FAR 27.917 Design. .. 530

422. FAR 27.921 Rotor Brake.. 531

423. FAR 27.923 Rotor Drive System and Control Mechanism Tests 532 423A. FAR 27.923 Rotor Drive System and Control Mechanism Tests 537 423B. FAR 27.923 Rotor Drive System and Control Mechanism Tests 538

424. FAR 27.927 Additional Tests.. 539 424A. FAR 27.927 Additional Tests.. 541

425.  FAR 27.931 Shafting Critical Speeds … 541

426. FAR 27.935 Shafting Joints... 546

427. FAR 27.939 Turbine Engine Operating Characteristics 546

**428.-445. RESERVED SECTION 26, FUEL SYSTEM**

446. FAR 27.951 General 567

447. FAR 27.952 Fuel System Crash Resistance 569

448. FAR 27.953 Fuel System Independence. 595

449. FAR 27.954 Fuel System Lightning Protection 596

450. FAR 27.955 Fuel Flow... 597 450A. FAR 27.955 Fuel Flow. 601

**451. RESERVED**

452. FAR 27.959 Unusable Fuel Supply. 605

453. FAR 27.961 Fuel System Hot Weather Operation 605 453A. FAR 27.961 Fuel System Hot Weather Operation 607

454. FAR 27.963 Fuel Tanks: General.. . 607 454A. FAR 27.963 Fuel Tanks: General . 608 454B. FAR 27.963 Fuel Tanks: General . 608

455. FAR 27.965 Fuel Tank Tests . 609

456. FAR 27.967 Fuel Tank Installation.. . 610

457. FAR 27.969 Fuel Tank Expansion Space . 612 457A. FAR 27.969 Fuel Tank Expansion Space ..613

458. FAR 27.971 Fuel Tank Sump.... ..613 458A. FAR 27.971 Fuel Tank Sump.. . 614

459. FAR 27.973 Fuel Tank Filler Connection.. .. 615 459A. FAR 27.973 Fuel Tank Filler Connection ..615

460. FAR 27.975 Fuel Tank Vents.. 616 460A. FAR 27.975 Fuel Tank Vents. .618 460B. FAR 27.975 Fuel Tank Vents. ..618

461. FAR 27.977 Fuel Tank Outlet. ..619

**462.-482. RESERVED**

### Page 11 (OCR)

SECTION 27, FUEL SYSTEM COMPONENTS Page No.

483. FAR 27.991 Fuel Pumps. .. 641 483A. FAR 27.991 Fuel Pumps.. 642

484. FAR 27.993 Fuel System Lines and Fittings 642

485. FAR 27.995 Fuel Valves. 643

486. FAR 27.997 Fuel Strainer or Filter.. .. 643 486A. FAR 27.997 Fuel Strainer or Filter. … 644

487. FAR 27.999 Fuel System Drains.. ..645 487A. FAR 27.999 Fuel System Drains.. 646

488.-497. RESERVED SECTION 28. OIL ŞYSTEM

498. FAR 27.1011 General 657

499. FAR 27.1013 Oil Tanks. 657

500. FAR 27.1015 Oil Tank Tests. 658

501. FAR 27.1017 Oil Lines and Fittings.. 659

502. FAR 27.1019 Oil Strainer or Filter 659 502A. FAR 27.1019 Oil Strainer or Filter 660

503. FAR 27.1021 Oil System Drains.. 661

504. FAR 27.1027 Transmission and Gearboxes: General 661

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

516. FAR 27.1041 General ..674 516A. FAR 27.1041 General 676

517. FAR 27.1043 Cooling Tests. 676

518. FAR 27.1045 Cooling Test Procedures 679 518A. FAR 27.1045 Cooling Test Procedures 680

**519.-530. RESERVED SECTION 30. INDUCTION SYSTEM**

531. FAR 27.1091 Air Induction.. .. 693 531A. FAR 27.1091 Air Induction.. .. 694

532. FAR 27.1093 Induction System Icing Protection.. .. 694 532A. FAR 27.1093 Induction System Icing Protection .706

**533.-547. RESERVED XI**

### Page 12 (OCR)

SECTION 31. EXHAUST SYSTEM Page No.

548. FAR 27.1121 General.. 722

549. FAR 27.1123 Exhaust Piping .. 723

**550.-560. RESERVED SECTION 32. POWERPLANT CONTROLS AND ACCESSORIES**

561. FAR 27.1141 Powerplant Controls: General 737 561A. FAR 27.1141 Powerplant Controls: General 738

**562. RESERVED**

563. FAR 27.1143 Engine Controls. 740 563A. FAR 27.1143 Engine Controls. 740 563B. FAR 27.1143 Engine Controls. 741

564. FAR 27.1145 Ignition Switches 741

565. FAR 27.1147 Mixture Controls 742

**566.-567. RESERVED**

568. FAR 27.1151 Rotor Brake Controls .745

569. FAR 27.1163 Powerplant Accessories 746

**570.-583. RESERVED SECTION 33. POWERPLANT FIRE PROTECTION**

**584. RESERVED**

585. FAR 27.1183 Lines, Fittings, and Components. 764

586. FAR 27.1185 Flammable Fluids. 764

587. FAR 27.1187 Ventilation 766

588. FAR 27.1189 Shutoff Means. 767

589. FAR 27.1191 Firewalls 768

590. FAR 27.1193 Cowling and Engine Compartment Covering. 772 590A. FAR 27.1193 Cowling and Engine Compartment Covering 773

591. FAR 27.1194 Other Surtaces. 774

592. FAR 27.1195 Fire Detector Systems... .775

**593.-616. RESERVED SECTION 34. EQUIPMENT - GENERAL**

617. FAR 27.1301 Function and Installation .. 804

618. FAR 27.1303 Flight and Navigation Instruments.. ..804

619. FAR 27.1305 Powerplant Instruments. ..805 619A. FAR 27.1305 Powerplant Instruments.. .. 806 619B. FAR 27.1305 Powerplant Instruments.. .. 806

620. FAR 27.1307 Miscellaneous Equipment ..807 XII

### Page 13 (OCR)

SECTION 34. EQUIPMENT - GENERAL (continued) Page No.

621. FAR 27.1309 Equipment, Systems, and Installations 808

622.-631. RESERVED SECTION 35. INSTRUMENTS - INSTALLATION FAR 27.1321 Arrangement and Visibility. .. 836

633. FAR 27.1322 Warning, Caution, and Advisory Lights. 836 FAR 27.1323 Airspeed Indicating System .. 839

635. FAR 27.1325 Static Pressure Systems 840

636. FAR 27.1327 Magnetic Direction Indicator 841

637. FAR 27.1329 Automatic Pilot System 842

**638.-639. RESERVED**

640. FAR 27.1335 Flight Director Systems. ... 852

641. FAR 27.1337 Powerplant Instruments.. 853 641A. FAR 27.1337 Powerplant Instruments.. 854

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

652. FAR 27.1351 General. ... 866

653. FAR 27.1353 Storage Battery Design and Installation. 868

**654. RESERVED**

655. FAR 27.1357 Circuit Protective Devices .. 871

**656.-657. RESERVED**

658. FAR 27.1361 Master Switch .. 875

659. FAR 27.1365 Electric Cables ..875

660. FAR 27.1367 Switches. 877

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

668. FAR 27.1381 Instrument Lights.. .. 885

669. FAR 27.1383 Landing Lights.. .. 885

670. FAR 27.1385 Position Light System Installation.. 886

671. FAR 27.1387 Position Light System Dihedral Angles.. . 886

672. FAR 27.1389 Position Light Distribution and Intensities. .. 886

673. FAR 27.1391 Minimum Intensities in the Horizontal.. .. 886 Plane of Forward and Rear Position Lights

674. FAR 27.1393 Minimum Intensities in Any Vertical.. .. 886 Plane of Forward and Rear Position Lights XIII

### Page 14 (OCR)

SECTION 37. LIGHTS (continued) Page No.

675. FAR 27.1395 Maximum Intensities In Overlapping. 886 Beams of Forward and Rear Position Lights

676. FAR 27.1397 Color Specifications. 886

677. FAR 27.1399 Riding Light.. . 886

678. FAR 27.1401 Anti-collision Light System 888

**679.-688. RESERVED SECTION 38. SAFETY EQUIPMENT**

689. FAR 27.1411 General.. 899

690. FAR 27.1413 Safety Belts. 901

691. FAR 27.1415 Ditching Equipment 902

RESERVED FAR 27.1419 Ice Protection 905

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

**702.-703. RESERVED**

704. FAR 27.1435 Hydraulic Systems 933

**705. RESERVED**

706. FAR 27.1457 Cockpit Voice Recorder 937

707. FAR 27.1459 Flight Recorders 938

708. FAR 27.1461 Equipment Containing High Energy Rotors 941

**709.-717. RESERVED SECTION 40. OPERATING LIMITATIONS**

718. FAR 27.1501 General.. 951

719. FAR 27.1503 Airspeed Limitations: General. 951

720. FAR 27.1505 Never-Exceed Speed ..... . 954

721. FAR 27.1509 Rotor Speed.. 956

**722. RESERVED**

723. FAR 27.1519 Weight and Center of Gravity .. 960

724. FAR 27.1521 Powerplant Limitations .. . 960 724A. FAR 27.1521 Powerplant Limitations ... 961 724B. FAR 27.1521 Powerplant Limitations .. . 962

**725. RESERVED**

726. FAR 27.1523 Minimum Flightcrew ...... . 965

727. FAR 27.1525 Kinds of Operations … 969

728. FAR 27.1527 Maximum Operating Altitude 969 XIV

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SECTION 40. OPERATING LIMITATIONS (continued) Page No.

729. FAR 27.1529 Instructions for Continued Airworthiness .. 970 (Maintenance Manual) 729A. FAR 27.1529 Instructions for Continued Airworthiness .. 972 (Maintenance Manual)

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

740. FAR 27.1541 General. 983

741. FAR 27.1543 Instrument Markings: General 983

742. FAR 27.1545 Airspeed Indicator .. 983

743. FAR 27.1547 Magnetic Direction Indicator 983

744. FAR 27.1549 Powerplant Instruments. 983

745. FAR 27.1551 Oil Quantity Indicators .. 983

746. FAR 27.1553 Fuel Quantity Indicator .. 984 747 FAR 27.1555 Control Markings. 985

748. FAR 27.1557 Miscellaneous Markings and Placards. 986

749. FAR 27.1559 Limitations Placard.. 987

750. FAR 27.1561 Safety Equipment 989

751. FAR 27.1565 Tail Rotor.. 989

**752.-761. RESERVED SECTION 42. ROTORCRAFT FLIGHT MANUAL**

762. FAR 27.1581 General. 1001

763. FAR 27.1583 Operating Limitations. 1009

764. FAR 27.1585 Operating Procedures ... 1010 764A. FAR 27.1585 Operating Procedures ... 1013

765. FAR 27.1587 Performance Information. 1013

766. FAR 27.1589 Loading Information 1015

**767.-774. RESERVED CHAPTER 3 - MISCELLANEOUS AIRWORTHINESS STANDARDS**

775. § 27 Appendix B Engineering Flight Test Guide for. 1025 Normal Category Helicopters - Instrument Flight Rules

776. Certification Procedure for Rotorcraft Avionics Equipment 1046 Standardized Test Procedure for Rotorcraft DC 1062 Electrical System Tests Standardized Test Procedure for Rotorcraft Generator Cooling 1068 Annunciator Panels. .. 1070 XV

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АС 27-1A 7/30/97 Chapter 3 (continued) Page No.

780. Definition of Engine Isolation Features as Applied to.. 1072 §§ 27.79(b)(2), 27.141 (b)(1), and 27.143(d)(1)

781. FAR 27.1541, 27.1543, 27.1545, and 27.1549 Instrument.. 1077 Markings 781A. FAR 27.1549 Powerplant Markings.. 1081 781B. FAR 27.1549 Powerplant Markings.. 1081

**782.-783. RESERVED**

784. Full Authority Digital Electronic Controls (FADEC) for .. 1084 Installations with Category A Engine Isolation

785. Agricultural Dispensing Equipment Installation 1086

786. Emergency Medical Service (EMS) Systems, Installations,. 1092 Interior Arrangements, and Equipment

787. Hour Meters or Time-in-Service Recording Devices in Rotorcraft .1110

788. Substantiation of Composite Rotorcraft Structure 1112 Appendix 1 Rotorcraft One-Engine-Inoperative (OEI) Power Assurance. A1-1 XVI

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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.

Figure/Table No. Title Page 58-1 H-V Requirements 71 58-2 IGE Controllability Requirements 72 58-3 Shaft Horsepower vs. Turbine Outlet 73 Temperature - Sea Level Standard Day 58-4 Uninstalled Takeoff Power Available 74 58-5 Installed Takeoff Power Available, Anti-Ice, 75 400 RPM 58-6 Power Assurance Check Chart 76 65-1 Autorotational Characteristics - Typical 87 69-1 Height-Velocity Diagram 101 69-2 Altitude/Weight Accountability 102 86-1 Static Longitudinal Stability 127 110-1 Demonstration Points 157 137-1 Load Factor Gross Weight Curve 184 140-1 Sample Yaw/Forward Speed Diagram 189 160-1 View Looking Forward 214 160-2 View Looking Forward 215 230-1 Comparison Percent of Time in Level Flight 287 XVII

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Figure/Table No. Title Page 250-1 Example of Casting Load Sheet Retract 316 Actuator Support-Landing Gear 250-1 Table 1: Castings to be Static Tested for Model XYZ 318 331-1 Cockpit Visibility 409 338-1 Sea State Code - World Meteorological 431 Organization 425-1 Cantilevered Shaft, First Critical Speed 545 425-2 Shaft Between Support Bearings, First Critical 545 Speed 447-1 Static Tension and Shear Loading Modes 589 447-2 Hose Assembly Tests 590 447-3 Minimum Average and Individual Loads for Hose 591 and Hose-End Fitting Combinations 447-4 Typical Method of Breakaway Fuel Fitting Load 592 Calculations 447-5 Typical Methods of Frangible or Deformable 593 Attachment Load Calculations 447-6 Typical Methods of Frangible or Deformable 594 Attachment Load Calculations 450-1 Fuel Flow 600 516-1 Additional Cooling Test Point 675 589-1 Table of Materials and Gages Acceptable for 771 Fireproof Protective Devices with Flat Surface Geometries 637-1 Deviation Profile 848 637-2 Operational Limitation 849 XVIII

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Figure/Table No. Title Page 693-1 Continuous Icing-Temperature vs. Altitude Limits 919 693-2 Intermittent Icing-Temperature vs. Altitude Limits 920 693-3 Continuous Icing-Liquid Water Content vs Drop 921 Diameter 693-4 Intermittent Icing-Liquid Water Content vs. Drop 922 Diameter 775-1 Rotorcraft Dynamic Stability Requirements for IFR 1045 780-1 Engine Isolation Rules 1076 785-1 Acceptable Ultimate Load Factor for Agricultural 1088 Dispensing Equipment Design 785-2 Sketch of Tank Pressure Test 1091 786-1 Typical Liquid Oxygen System 1108 786-2 Typical Liquid Oxygen System - Using 1109 Combination Valve APX1-1 Installed Engine Flow Capacity Check A1-9 XIX

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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 FPM feet per minute FPS feet per second GPS global positioning system HF high frequency HIRF high intensity radiated fields HRD high rate of discharge HUMS health and usage monitoring systems HV height-velocity IAS indicated airspeed ICAO International Civil Aviation Organization ICS inter-communication system IFR instrument flight rules GE 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 XX

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7/30/97 OAT outside air temperature OEl one engine inoperative OGE out of ground effect PBA pitch bias actuator 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 AC 27-1A WAT weight, altitude, temperature Altitudes HD density altitude HP pressure altitude V speeds VD diving speed VH speed in level flight with maximum continuous power VMO maximum operating limit speed VNE never-exceed speed VTOSS takeoff safety speed for Category A rotorcraft VX speed for best angle of climb VY speed for best rate of climb MMO maximum operating mach number N speeds NF free turbine speed NG gas generator speed NP power turbine speed NR rotor speed Coefficients CD coefficient of drag CL coefficient of lift CP coefficient of power CT coefficient of thrust XXI

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**CHAPTER 1. PART 21**

CERTIFICATION PROCEDURES FOR PRODUCTS AND PARTS (Amendment 21-50)

**1.-3. RESERVED.**

4. § 21.16 SPECIAL CONDITIONS.

a. Ihe 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-110. That office will ensure coordination with the affected agency and industry elements including the Assistant Chief Counsel.

All comments will be considered and the disposition documented by the Rotorcraft Directorate. ASW-100 will issue the special conditions.

b. Basis for Development.

(1) Special conditions are justified on the basis of the existing Part 27 being inadequate or inappropriate due to novel or unusual design features of the rotorcraft to be certificated.

(2) The phrase "novel or unusual" as used in § 21.16 is a very relative term. As used hereafter in applying § 21.16 to justify the issuance of special conditions, "novel or unusual will be taken with respect to the state of technology envisaged by the applicable airworthiness standards of this subchapter. It must be recognized that in some areas which will vary from time to time, the state of the regulations may somewhat lag the state of the art in new design because of the rapidity in which the state of the art is advancing in civil aeronautical design and because of the time required to develop the experience base needed by the FAA/AUTHORITY to proceed with general rulemaking. Applicants for type certification of a new design have the opportunity to mitigate the impact of not knowing the precise airworthiness standards to be applied for "novel or unusual design features" by consulting with the FAA/AUTHORITY early in their certification planning when such features are suspected or known by the applicant to exist. It should also be recognized that, because of the intentional objective nature of the airworthiness standards of this subchapter, many new design features which might be thought of as "novel or unusual" may already be adequately covered by existing regulations, thus obviating the need to issue special conditions.

Par 1 1

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(3) Before proposing special conditions, the certification staff should very thoroughly analyze the existing regulations and ensure they are inadequate or inappropriate in light of a new and novel design feature.

Par 4

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**5.-7. RESERVED.**

Intentionally Left Blank Par 5 3 (thru 5)

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8. § 21.31 TYPE DESIGN.

The regulatory basis for requiring data to define the design is contained in § 21.31.

This section is self-explanatory and broad enough in scope to give the certification staff access to sufficient data to determine compliance with Part 27.

Par 8

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9.-11. RESERVED Intentionally Left Blank Par 9 7 (thru 9)

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АС 27-1A 7130/97

12. § 21.33 INSPECTION AND TESTS.

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

(1) Ensure the test rotorcraft conforms to the type design. This must be accomplished prior to presentation to the FAA/AUTHORITY for testing.

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

b. FAA/AUTHORITY Responsibility.

(1) The design evaluation engineers should ensure that the type design is adequate in their technical area and that the inspections and tests to be conducted are appropriate and sufficient to show compliance with Part 27.

(2) As changes to the rotorcraft are made during the test program, the flight test crew should ensure that the appropriate design evaluation engineer concurs with the change and the conformity inspection of the change has been conducted.

10 Par 12

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AC 27-1A 7/30/97

**13.-15. RESERVED.**

Intentionally Left Blank 11 (thru 13) Par 13

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AC 27-1A 7130/97

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

a. Explanation.

(1) This section outlines the requirements of the applicant for aircraft type certification and should be used in conjunction with 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 FAA/AUTHORITY in report form for evaluation to determine what verification flight tests the FAA/AUTHORITY may elect to conduct. The report should conclude that in the applicant's opinion the test aircraft complies with the applicable certification requirements. The FAA/AUTHORITY verification flight test should include, but not be limited to, the critical or marginal results contained in the applicant's flight test report.

The FAA/AUTHORITY's role in the certification effort is not envisioned to be one of conducting day-to-day routine flight tests with the applicant, but only to verify his results through limited sampling. In certain tests, such as high altitude testing at a remote mountain site, there is an advantage in conducting flight tests concurrently with the applicant. Additionally, the FAA/AUTHORITY can provide technical flight test assistance to the applicant in certain cases. This can be done after a cursory review and a letter of authorization is issued to the flight test crew.

(2) Preflight Test Planning. After the applicant's flight test report is reviewed, it should be determined what FAA/AUTHORITY engineering flight tests are necessary.

These tests are normally specified in the Type Inspection Authorization (TIA). At the same time the FAA/AUTHORITY must know and agree to the applicant's proposed means of data acquisition, reduction, and expansion of the flight test data. The adequacy of the test instrumentation should be evaluated prior to official type certification tests (reference Paragraph 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.

(ill) Engine cooling.

(4) Test Groupings.

14 Par 16

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Weight and CG. In addition to the regulatory relationship of one test to another, efficient testing requires that consideration be given to the accomplishment of as many tests on a single flight as can be accommodated successfully.

(ii) Special Instrumentation. Similarly, consideration should be given to grouping of tests that involve special instrumentation. Examples of these are takeoff and landing tests which usually require group equipment to record horizontal distance, height, and time. Ground calibration of the airspeed indicating system can be accomplished at the same time. It is the applicant's responsibility to provide the necessary instrumentation.

b. Procedures.

(1) Type Certification Flight Tests.

(i) Prior to initiating official FAA/AUTHORITY flight tests, a conformity inspection of the test aircraft must be accomplished. This is needed to ensure that the test aircraft is in the proper configuration or "conforms" to the engineering drawings and documents that have been submitted to FAA/AUTHORITY, evaluated, and approved. It is absolutely essential to know the configuration being tested in any engineering flight evaluation. Conformity inspection prior to TIA flight tests assures that testing will not be wasted because of configuration uncertainties.

(ії) Certification Handbook 8110.4, Paragraph 67, contains a requirement that the applicant must keep the FAA/AUTHORITY advised of any configuration changes to the aircraft. The manufacturing inspector should keep the FAA/AUTHORITY flight test pilot apprised of any change which may affect safety of the test aircraft or may influence test results.

(ili) Results of the conformity inspection and the engineering flight test program must be documented. This is normally done in the Type Inspection Report (TIR). Results may be documented in any acceptable engineering format. The report should be in sufficient detail to clearly show how compliance with each appropriate section of the rule was determined.

(iv) The flight test pilot must ensure that the FAA/AUTHORITY manufacturing inspector and the certification engineer are aware of all configuration changes found necessary as a result of FAA/AUTHORITY tests. The manufacturing inspector is responsible for ensuring that all changes are incorporated into production drawings after the design data reflecting the change have been approved by the certification engineer.

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(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.

(1) A comprehensive and systematic check of all aircraft components must be made to ensure that they perform their intended function and are reliable.

(її) 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(1) above to ensure that significant changes resulting from type certification tests can be incorporated on the aircraft prior to F&R tests.

(ili) 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.

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17.-23. RESERVED Intentionally Left Blank Par 17 17 (thru 23)

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24. § 21.39 (Amendment 21-59) FLIGHT TEST INSTRUMENT CALIBRATION AND CORRECTION REPORT.

a. Explanation. It is the applicant's responsibility to provide instrumentation for all parameters needed to show compliance with the airworthiness regulations.

(1) For those data which are necessary to show compliance with the regulations, a permanent record should be established. A permanent record is acceptable in either graphical or photographic form, and in some instances a manual recording may be satisfactory.

(2) Regardless of the record form, the accuracy of the record must be established by reference to a laboratory standard traceable to the National Bureau of Standards.

(3) If multiplexing is used, the time base must be synchronized to a reference point from which the magnitude of each parameter can unquestionably be determined.

Also, the sampling rate should be sufficiently frequent to ensure that the maximums, minimums, and trends of magnitude of the parameter are recorded with respect to time.

b. Procedures. Prior to conducting flight tests, the FAA/AUTHORITY flight test team should review the applicant's flight test instrumentation, calibration, and correction report.

(1) The frequency of recalibration varies with the consistency of the instrumentation under consideration. For example, cyclic and collective position is sometimes calibrated immediately before and after a flight where these parameters are used to provide critical flight data. Six months is a typical interval for recording/signal conditioning and nonstrain gage sensors, while 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.

**25.-30. RESERVED.**

24 (thru 30) Par 24

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**CHAPTER 2. PART 27**

**AIRWORTHINESS STANDARDS**

**NORMAL CATEGORY ROTORCRAFT SECTION 1. GENERAL**

31. § 27.1 APPLICABILITY.

a. Explanation. This section prescribes the rotorcraft categories eligible for certification under this Part. There is no minimum weight limit for certification under Part 29; however, Part 27 is applicable to rotorcraft with maximum weights of 6,000 pounds or less.

(1) Without Engine Isolation. For single-engine rotorcraft and multiengine rotorcraft without engine isolation, the height-velocity (HV) diagram is conducted with sudden failure of all engines, and the takeoff maneuver must pass through the clear area of the diagram to the 50-foot point with all engines operating.

(2) With Engine Isolation. Part 27 multiengine rotorcraft may be certificated with engine isolation features (reference Paragraph 780 of this AC). These rotorcraft are not required to meet the Part 29, Category A, performance requirements, and continued flight after an engine failure is not assured since under some conditions tallure of the remaining engine may occur after a limited time. The takeoff is conducted with all engines operating, while the height-velocity diagram is determined with the most critical engine inoperative. If complete Part 29, Category A, design features and performance are achieved, the Category A performance may be included in the FAA/AUTHORITY-approved portion of the Rotorcraft Flight Manual although this performance is not required by the regulations.

b. Procedures. None.

31A. § 27.1 (Amendment 27-33) APPLICABILITY.

a. Explanation. Amendment 33 formally introduced the requirements for certification of a Part 27 rotorcraft to Category A design and performance standards.

These standards are found in Appendix C of Part 27. The establishment of Category A design and performance for multiengine Part 27 rotorcraft is still voluntary. If so requested, the corresponding AC 29-2A material applies.

b. Procedures. None

32. § 27.2 (Amendment 27-28) SPECIAL RETROACTIVE REQUIREMENTS.

a. Explanation.

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(1) Amendment 27-28 requires a combined shoulder harness and safety belt (also called a torso restraint system) at each occupant's seat for all rotorcraft manufactured after September 16, 1992.

(2) The design features of the restraint system are mainly contained in this section rather than having to refer to other sections within Part 27 except for a general reference to the differing strength standards between earlier static strength only standards and the static and dynamic strength standards of Amendment 27-25.

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

Those rotorcraft type designs certificated to static strength standards alone prior to Amendment 27-25, such as 4 g's forward may use belt and harness systems, characterized as 1,500 pounds strength systems, provided they comply with those standards. TSO C22f and earlier restraint systems have such ratings. A combined belt and harness with a 1,500 pounds rating, which comply with the Part 27 standards for the rotorcraft type design, but are not necessarily TSO approved, may be approved as a part of the type design. Such design information for a non-TSO'd item would be included in a note on the aircraft type certificate data sheet (TCDS) or specification sheet by part number as "required equipment." TSO C114-approved torso restraint systems, characterized as 3,000 pounds strength system, may be used provided the design features comply with this section, but no special information on the TCDS is necessary Those rotorcraft type designs certified to dynamic test requirements of Amendment 27-25 should use torso restraint systems approved under TSO C114 or approved under equivalent standards such as those contained in Part 27.

(4) Load Distribution and Design Requirements. Although not stated in § 27.2, a 60 percent and 40 percent load distribution between the safety belt and harness, respectively, is required in § 27.785(g). The safety belt should withstand 100 percent if the safety belt is capable of being used alone. Also, the safety belt or harness attachments to the seat or structure should include the 1.33 factor described in § 27.785(f)(2) of Amendment 27-21 for those rotorcraft with that certification criteria or should include the 1.15 factor as described in § 27.625 (and predecessor Part 6) standards for those rotorcraft with the earlier certification criteria. A factor is used whether test or analysis methods are used for static substantiation of the seating systems. Refer to Paragraph 336b, (§ 27.785), of this AC.

(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.

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b. Procedures.

(1) A TSO-approved combined safety belt and harness or torso restraint system may be used provided the installation requirements in § 27.2 are satisfied. A combined belt and harness (not necessarily TSO approved) may be approved as a part of the rotorcraft type design and so noted on the aircraft specification or TCDS.

(2) Structural analysis or static test may be used. For those rotorcraft designs that are subject to the dynamic test standards of § 27.562, the torso restraint system is required to be qualified for the particular use or installation in each rotorcraft type design. A dynamic test may be required for alternate restraint systems as well as the originally approved system. TSO C114 approval does not constitute approval for installation of a restraint system in a rotorcraft design subject to dynamic test.

AC 20-137 dated March 30, 1992, concerns in part the dynamic test standards of Amendment 27-25.

(ii) AC 23-4 dated June 20, 1986, concerns static test procedures for small airplane seats and restraint systems. (Certain small airplanes manufactured after December 12, 1986, should have harnesses for each seat also.) A test proposal for rotorcraft installations may adopt procedures appropriate to the particular installation.

The 60/40 percent distribution is sufficiently achieved when the blocks in Figure 4 of AC 23-4 are used.

(iii) The static design side load for the harness installation may be proven by test or analysis using the load distribution previously noted. For "older" designs, the side load of § 27.561(b)(3) (iii) is 2.0g, and for later designs (Amendment 27-25 and later), it is 8.0g.

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**33.-41. RESERVED.**

Intentionally Left Blank Par 33 34 (thru 42)

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**SECTION 2. FLIGHT - GENERAL**

42. § 27.21 (Amendment 27-21) PROOF OF COMPLIANCE.

a. Explanation.

(1) This section provides a degree of latitude for the FAA/AUTHORITY test team in selecting the combination of tests or inspections required to demonstrate compliance with the regulations. Compliance should be shown for applicable combinations of gross weight, center of gravity, altitude, temperature, airspeed, rotor RPM, etc. Engineering tests are designed to investigate the overall capabilities and characteristics of the rotorcraft throughout its operational envelope. Testing will identify operating limitations, normal and emergency procedures, and performance information to be included in the FAA/AUTHORITY-approved portion of the flight manual. The testing must also provide a means of verifying that the rotorcraft's actual performance, structural design parameters, propulsion components, and systems operations are consistent with all certification requirements.

(2) Section 21.35 requires, in part, that the applicant show compliance with the applicable certification requirements, including flight test, prior to official FAA Type Inspection Authorization (TIA) testing. Compliance in most cases requires systematic flight testing by the applicant. After the applicant has submitted sufficient data to the FAAJAUTHORITY showing that compliance has been met, the FAA/AUTHORITY will conduct any inspections, flight, or ground tests required to verify the applicant's test results. FAA/AUTHORITY compliance may be partially determined from tests conducted by the applicant if the configuration (conformity) of the rotorcraft can be verified. Compliance may be based on the applicant's engineering data and a spot check or validation through FAA/AUTHORITY flight tests. The FAA/AUTHORITY testing should obtain validation at critical combinations of proposed flight variables if compliance cannot be inferred using engineering judgment from the combinations investigated.

(3) Performance tests include minimum operating speed (hover), takeoff and landing, climb, glide, height-velocity, and power available. Certain other performance tests, such as critical engine survey for multiengine installations, may be conducted to meet specific requirements. Detailed performance test procedures and allowable extrapolation or simulation limits are contained in the respective Paragraphs in this AC.

(i) Hover tests are conducted to determine various combinations of altitude, temperature, and gross weight for both in-ground-effect (IGE) and, if required by the applicant, out-of-ground effect (OGE) conditions. From these data, the hover ceiling may be calculated.

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(її) Takeoff and landing tests are conducted to determine that a takeoff or landing can be safely executed without requiring exceptional piloting skill or favorable conditions at any approved combination of altitude, temperature, and gross weight.

(ili) For rotorcraft other than helicopters, climb tests establish the variations of rate-of-climb at the best rate-of-climb or published climb airspeed(s) at various combinations of altitude, temperature, and gross weight. For helicopter, climb tests are conducted as required to determine the best rate-of-climb speed, Vy.

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

(V) Power available tests are conducted to verify the calculated installed specification engine performance model on which published performance is based.

(4) The purpose of rotorcraft stability and control tests is to verify that the rotorcraft possesses the minimum qualitative and quantitative flying qualities and handling characteristics required by the applicable regulations. In order to assess the handling qualities, standardized test procedures must be utilized and the results analyzed by accepted methods. Section 27.21(a) allows calculation and inference which includes extrapolation and simulation, whereas § 27.21(b) requires demonstration of controllability, stability, and trim. Combinations of § 27.21(a) and (b) may be used to show compliance with the operating envelope limits. Test methods and equipment are described in individual paragraphs of this advisory circular.

b. Procedures.

(1) Efforts should begin early in the certification program to provide advice and assistance to the applicant to ensure coverage of all certification requirements. The applicant should develop a comprehensive test plan which includes the required instrumentation.

(2) The tests and findings specified in Paragraph 42a(3) are required of the applicant to show basic airworthiness and probable compliance with the minimum requirements specified in the applicable regulations. After these basic findings have been submitted and reviewed, a Type Inspection Authorization, or equivalent, can be issued. The FAA/AUTHORITY will develop a systematic plan to spotcheck and confirm that compliance with the regulations has been shown. The test plan will consider combinations of weight, center of gravity, and RPM and cover the range of altitude and temperature for which certification is requested.

43. § 27.25 (Amendment 27-14) WEIGHT LIMITS.

a. Explanation.

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(1) This section is definitive and specifies criteria for establishing maximum and minimum certificating weights. These weights may be based on those selected by the applicant, design requirements, or the limits for which compliance with all applicable flight requirements has been shown.

(2) Typical requirements that may establish the maximum and minimum weight limits include:

(i) Maximum: Structural limits, performance requirements, stability, and controllability requirements.

(II) Minimum: Autorotative rotor RPM, stability, and controllability requirements.

(3) Jettisonable External Cargo.

(i) Section 27.25(c) was added by Amendment 27-11 to provide a basis for approving an increased gross weight that would be an external jettisonable load.

Section 27.865, "External load attaching means, " includes hoist and hook design features for the load attaching devices that were added to Part 27 but removed from § 133.43. Part 133, "Rotorcraft External-Load Operations," was also amended (Amendment 133-5) concurrently to complement the changes to Parts 27 and 29.

(Ш) Approvals under the policy in Review Cases Nos. 37 and 55 of FAA Order 8110.6 were no longer necessary. These review cases concerned the policy/standards for external cargo configurations using a cargo hook whenever the standard limitations were exceeded. If the standard limitations were not exceeded, external cargo hooks and hoists and external cargo configuration approvals could be made under Part 133, Subpart D, prior to Amendment 133-5.

In the preamble of Amendment 27-11 (Proposal 2-99, 41 FR 55454;

December 20, 1976) the agency stated that"...§ 27.25(c) and § 29.25(c) are intended to provide only a total weight standard for approving the rotorcraft structure for operation under Part 133." The policy in Review Case No. 55 also indicates the powerplant or propulsion system is also subject to evaluation for the increased weight.

As indicated in § 27.865, fatigue substantiation of the external cargo attaching means is not required. The rotorcraft structure, rotors, etc., are only subject to fatigue evaluation under § 27.571 whenever the standard structural limitations are exceeded (Review Case No. 55).

(iv) Whether or not the standard limitations are exceeded, the flight characteristics evaluations/standards of § 133.41 are appropriate even for engineering approval. Section 133.41 is also appropriate for the individual operator to obtain his operating certificate. The operator may use an FAA/AUTHORITY approved RFM Par 43 45

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supplement to prepare his own rotorcraft load combination flight manual required by § 133.47.

b. Procedures.

(1) It may not be possible to demonstrate quantitatively all the flight requirements at the minimum weight because of test instrumentation requirements.

The test team must ensure that the rotorcraft complies with the applicable requirements at the lowest permissible flying weight. This evaluation may be done qualitatively with the test instrumentation removed and with minimum crewmembers if no critical areas exist or are anticipated. Additionally, reasonable extrapolation is permitted. However, if critical areas at minimum flying weights are apparent, extrapolation should not be permitted.

(2) Whenever a gross weight increase under § 27.25(c) is requested, a TIA evaluation is necessary to evaluate the new limitations and ensure that § 133.41 for typical or representative cargo weights and/or shapes (or density) is satisfactory. All possible combinations of weights and shapes are not evaluated. The representative configurations may be noted in the RFM or RFM Supplement for the operator's information. Sections 133.41 and 133.47 must be satisfied by the individual operator for the particular case at hand. The approved RFM or RFM Supplement should provide the necessary limitations and any other information about the representative cargo configurations evaluated. Section 133.41 also permits the operator to obtain approval of additional and unique cargo configurations provided approved limitations are observed. Paragraph 762 of this AC concerns the RFM and its contents.

(3) See AC 29-2, Certification of Transport Category Rotorcraft Paragraph 230, concerning § 29.571, for fatigue substantiation and external cargo considerations that apply to § 27.571 as well.

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

44. § 27.27 (Amendment 27-2) CENTER OF GRAVITY LIMITS.

a. Explanation

(1) This regulation is definitive and requires that the center of gravity limits be defined. Proof of compliance with all applicable flight requirements is required within the range of established CG's. Along with the longitudinal CG limits, the lateral CG limits should either be established or determined to be not critical.

(2) Ballast is usually carried during the flight test program to investigate the approved gross weight/center of gravity limits. Lead is the most commonly used form of Par 43 46

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ballast during rotorcraft flight testing although other types of ballast, such as water, may serve just as well. Water may have the added benefit of being jettisonable during critical flight test conditions. Care must be taken regarding the location of ballast. The strength of the supporting structures should be adequate to support such ballast during the flight loads that may be imposed during a particular test and for the ultimate inertia forces of § 29.561(b)(3). Of critical importance is the method of securing the ballast to the desired locations. To avoid any undesired in-flight movements of the ballast, a positive method of constraint is mandatory. The flight test crews should also visually verify the amount, location, and integrity of the ballast. The effects of mass moment of inertia on the flight characteristics due to the ballast locations should also be considered. The mass moment of inertia of the test rotorcraft should, to the extent possible, be the same as that expected in normal, approved loadings, especially during tests involving dynamic inputs.

b. Procedures.

(1) Center of gravity locations and limits are of prime importance to rotorcraft stability and safety in flight. The primary concern is establishment of the longitudinal center of gravity limits. Lateral center of gravity limits with respect to longitudinal center of gravity limits are also important. The design of the rotorcraft is usually such that approximate lateral symmetry exists. This lateral symmetry can be upset by numerous probable lateral loadings possibly resulting in the necessity to establish lateral center of gravity limits. Stability and control characteristics may be seriously affected by loading outside the established center of gravity limits. The established center of gravity limits must be that as fuel is consumed, it is possible for the rotorcraft to remain within the established limits by acceptable loading and/or operating instructions.

(2) Structural limits may restrict the maximum forward longitudinal center of gravity limits. However, in most cases it is the maximum value established wherein adequate low speed control power exists to meet such requirements as § 27.143(c).

Likewise, the maximum aft center of gravity limit may be a "structural limit," but it usually is determined during flight test after the rotorcraft's handling qualities tests have been conducted. Flight tests may reduce the "structural limit" CG envelope, but flight tests alone should not be used to expand the "structural limit." Additional items which may influence the maximum aft center of gravity limits may be malfunctions of automatic stabilization equipment, excessive rotorcraft attitudes during critical phases of flight, or adequate control power to compensate for an engine failure.

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

(4) In summary, it is of prime importance that longitudinal and lateral center of gravity limits be determined so that unsafe conditions do not exist within the approved Par 44 47

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altitude, airspeed, ambient temperature, gross weight, and rotor RPM ranges. All relevant malfunctions must be considered.

45. § 27.29 (Amendment 27-14) EMPTY WEIGHT AND CORRESPONDING CENTER OF GRAVITY.

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

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

(2) Compliance with Paragraph (b) of § 27.29 is accomplished by the use of an equipment list which defines the installed equipment at the time of weighing and the weight arm and moment of the equipment.

b. Procedures.

(1) Determination of the empty weight and corresponding center of gravity is primarily the responsibility of the manufacturing inspector. This determination is normally made on the production rotorcraft rather than the prototype. If the manufacturer wishes to avoid the necessity of weighing each production rotorcraft and he has been issued a production certificate, he may make a detailed proposal defining the procedures he will use to establish an empty weight and CG When his proposal is approved, he will weigh the first five to ten production rotorcraft and show that the rotorcraft will be within 11 percent on empty weight and t0.2 inches on CG After this procedure is established, the empty weight and CG may be computed except that at regular intervals, a rotorcraft will be weighed to ensure the tolerances are still being maintained; e.g., one in ten rotorcraft.

(2) For prototype and modified rotorcraft, it is only necessary to establish a known basic weight and CG position (by weighing) from which the extremes of weight and CG travel required by the test program may be calculated. See AC 91-23 (Pilots Weight and Balance Handbook) for a sample weight and balance procedure.

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

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

c. Ballast Loading and Type.

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

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

46. § 27.31 REMOVABLE BALLAST.

a. Explanation. This regulation provides the option of using removable ballast to obtain desired center of gravity locations to determine compliance with the flight requirement of this Part. Fixed ballast used for flight operations after type certification must be documented in the type design data. Removable ballast is used primarily on small rotorcraft to control the CG with different passenger loadings although this regulation does not permit its use on transport rotorcraft. If removable ballast is used, the rotorcraft flight manual must include instructions regarding its use and limitations.

b. Procedures. None

47. § 27.33 (Amendment 27-14) MAIN ROTOR SPEED AND PITCH LIMITS.

a. Explanation.

(1) General. This section requires the establishment of power-on and power-off main rotor speed limits and the requirements for low rotor speed warning.

(2) Power-On. The power-on limits should be sufficient to maintain the rotor speed within these limits during any appropriate maneuver expected to be encountered in normal operations throughout the flight envelope for which certification is requested.

In the past a minimum power-on range of approximately 3 percent has been required due to engine governor and engine operating characteristics. With the introduction of advanced engines and electronic engine controls, there may not be a need for a range.

One fixed value may suffice. If substantiated, transient power-on values may also be acceptable.

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(3) Power-Off. The power-off rotor speed limits should be sufficient to encompass the rotor speeds encountered during normal autorotative maneuvers except for final landing phase (touchdown) for which rotor RPM may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. The limits should also be sufficient to cover the ranges of airspeed, weight, and altitudes for which certification is requested. It is not the intent of the rule to require the minimum and maximum limit values in conjunction with extremes such as maximum/minimum weights and/or high altitude. The minimum and maximum rotor speed requirements should be thoroughly evaluated at normal operating environment; l.e., at altitudes from approximately sea level to 10,000 feet, temperatures not at extremes, and weights as necessary for other tests and as required to readily establish the limit rotor speeds.

Spot checks of the autorotative requirements should be made at the extremes of the flight envelope and environmental conditions during normal tests at those conditions.

Under conditions where high autorotative rotor speeds may be encountered, it is acceptable for the pilot to adjust the controls to prevent overspeeding of the rotor. At light weight combined with low altitudes and extremely cold temperatures, the normal low pitch setting may not be sufficient to maintain autorotational rotor speed values within limits. If this occurs, the manufacturer may elect to adjust the low pitch stops as a maintenance procedure at extreme ambient conditions provided the flight and maintenance manuals clearly present the rigging requirements and procedures. There must be sufficient "overlap" of ambient conditions between configurations such that rerigging is not required whenever ambient temperature and surface elevation change slightly. Any downrigging of the low pitch stop must continue to ensure adequate clearance between controls and other rotorcraft structure and should be evaluated during flight test. Both the power-on and power-off limits may also be established by encountering critical flapping limits in some approved flight conditions such as high airspeed or sideward tight.

(4) Low Speed Warning. If it is possible under expected operating conditions for the rotor speed to fall below the minimum approved values, the requirement exists for a low rotor speed warning. This warning is required on all single-engine rotorcraft and on multiengine rotorcraft where there is not an automatic increase in remaining engine(s) power output upon failure of an engine. Although not required by the rule, essentially all of today's multiengine rotorcraft have a low rotor speed warning system installed. If the minimum power-on and power-off rotor speed limits are different, the warning signal should be at the higher speed, normally the power-on minimum rotor speed. One type of rotorcraft has a warning system cutout if the collective is full down, and other types have other warnings on the engine speed to indicate engine failure. All of these related warning systems must be evaluated with emphasis on ensuring adequate rotor speed.

b. Determination and Testing. Refer to Paragraph 721 (§ 27.1509) for additional information on determining and testing rotor limits.

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**48.-56. RESERVED.**

Intentionally Left Blank Par 48 51 (thru 59)

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AC 27-1A 7/30/97/

**57. RESERVED.**

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**7130/97 AC 27-1A SECTION 3. PERFORMANCE**

58. § 27.45 (Amendment 27-21) GENERAL

a. Explanation.

(1) Section 45 of Part 27 lists some of the rules and standards under which the performance requirements are to be met. This paragraph will provide general guidelines that may be used throughout a flight test program. It is impossible to find ideal test conditions and there are many variables which affect the flight test results that must be taken into account. Some of these variables are wind, temperature, altitude, humidity, rotorcraft weight, power, rotor RPM, center of gravity, etc. A thorough knowledge of the testing procedures and data reduction methods is essential and good engineering judgment must be used to determine acceptable test conditions. The test results should be analyzed and expanded by approved methodology within the guidelines of this paragraph.

(2) Performance should be based on approved engine power as determined in Paragraph b(5) below and not on any transient limits. Approved transient limits are basically for inadvertent overshoots of approved operational limits and any sustained operation in these transient limit areas usually requires some form of special maintenance. However, for such demonstrations as landing procedure demonstration and height-velocity (HV) determination, low rotor speeds (within approved limits) have been authorized. Such transients, if authorized, must be flight evaluated.

(3) Where variations in the parameter on which a tolerance is allowed will have an appreciable effect on the test, the results should be corrected to the standard value of the parameter; otherwise, no correction is necessary.

b. Procedures.

(1) Winds for Testing.

Allowable wind conditions will vary with the type of test and will also be different for different types and gross weight rotorcraft. For example, higher winds can usually be tolerated for takeoff and landing tests than for hover performance.

Higher winds can sometimes be tolerated during hover performance testing on rotorcraft with high rotor downwash velocities. Generally, unless the effects of wind on hover performance tests can be determined and/or accounted for, hover performance testing should be conducted in winds of 3 knots or less.

(i¡) In-ground-effect controllability and maneuverability testing should be conducted in surface winds of less than 5 knots, or when higher steady wind conditions exist, with a maximum gust spread of 5 knots.

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АС 27-1A 7/30/97 As can be seen from the foregoing, there is no such thing as an exact allowable wind for a particular test or rotorcraft. The flight test team must decide on the allowable wind for each condition based on all available information and their engineering judgment. The following summary of allowable wind conditions is given for general guidance only:

(A) Hover performance - 0 to 3 knots.

(B) Height-velocity - 0 to 3 knots.

(C) IGE controllability and maneuverability - 0 to 5 knots.

(iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests.

(2) Altitude Effects. Using FAA/AUTHORITY-approved methodology, hover performance may be extrapolated and/or interpolated from test data up to a maximum of +4,000 feet. Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics should not be extrapolated higher than approximately 2,000 feet density altitude from the test altitude. Cruise stability/controllability tests should be evaluated at least at two different altitudes, the lowest practical altitude and approximately the highest cruise altitude requested for approval. This can allow an interpolation of approximately 10,000 feet. As in all testing, extrapolation and/or interpolation should only be considered if all available information and engineering judgment indicate that regulatory compliance can be met at the untested conditions.

(3) Altitude Limitations.

Explanation.

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

- Maximum operating altitude, and

- Maximum takeoff and landing altitude.

(B) Maximum operating altitude is an operating limitation required by § 27.1527 and delineates the maximum altitude to which operation is allowed. This altitude normally constitutes the maximum cruise or en route altitude.

(C) Maximum takeoff and landing altitude is the hover in-ground-effect (IGE) ceiling for a rotorcraft as described in § 27.73. The hover ceiling and any information pertinent to takeoff and landing are presented in the performance information section of the RFM. For rotorcraft certified to CAR 6, Amendment 6-7 or any amendment of FAR 27, a hover ceiling may not be presented above the altitude at which H-V and IGE controllability tests were conducted plus allowable extrapolation, Par 58 62

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unless that extrapolated altitude is at least 7,000 feet. If the applicant elects to demonstrate these tests to an altitude below 7,000 feet, then that altitude is the maximum takeoff and landing altitude of the rotorcraft. The maximum takeoff and landing altitude may be coincident with, but never above the maximum operating altitude limitation. Takeoff and landing and hover ceiling data and presentation requirements are presented in §$ 27.51, 27.73 and 27.1587.

(її) Procedures.

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

(1) Takeoff (§27.51)

(2) Climb (§§ 27.65 and 27.67)

(3) Performance at minimum operating speed (§ 27.73)

(4) Landing (§ 27.75)

(2) Limiting height-speed envelope (§ 27.79)

(6) GE controllability (§ 27.143c) (Z) Cooling (§§ 27.1041, 27.1043 and 27.1045)

(8) Engine operating characteristics (§ 27.939) Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this AC.

(B) As detailed in subparagraph b(2) above, the maximum allowable extrapolation of H-V, IGE controllability and engine operating characteristics is +2,000 feet. Therefore, the maximum takeoff and landing altitude presented in the RFM is not normally more than 2,000 feet above the density altitude experienced at the high altitude test site, or for CAR 6, Amendment 6-7 and subsequent, unless test results were demonstrated to at least 7,000 feet.

(C) If IGE controllability is demonstrated to at least 17 knots of wind at 7,000 feet, hover capability above 7,000 feet may be presented provided that the maximum demonstrated safe wind for takeoff and landing above 7,000 feet is specified in the RFM.

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АС 27-1А 7/30/97 (D) The requirements for data collection and presentation in the RFM vary depending upon the certification basis of the rotorcraft. These requirements are presented by regulation and amendment in Figures 58-1 and 58-2.

(E) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in Paragraph b(2) and not above the lowest limiting altitude resulting from the requirements of subparagraph A of this paragraph.

(4) Temperature Effects.

Background.

(A) In the past, approved analyses were frequently accepted for determining the extreme temperature effects on performance and flight characteristics.

With the introduction of newer, higher performance rotorcraft, advanced rotor blade designs, higher airspeeds, and higher blade tip Mach numbers, the previous methods have proven to be insufficient. Therefore, the performance and flight characteristics should be validated at extreme temperatures; however, analysis may be permitted if a suitable methodology is demonstrated.

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

(C) Additionally, design deficiencies surfaced when the rotorcraft were exposed to temperature extremes and some of these difficulties were severe enough to require the redesign of equipment and/or materials. Therefore, to satisfy § 27.1309(a), the applicant needs to substantiate the total rotorcraft throughout the foreseeable range of operating temperatures.

(її) Procedures.

(A) The FAA/AUTHORITY is responsible for verifying the effects of temperature on performance and handling characteristics. A limited flight verification, if necessary, could include spot checks of hover performance, IGE controllability, vibration, simulated power failure, static stability, height-velocity, VNE/No evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.

(B) Extrapolation of test data should only be allowed if the applicant's predicted or calculated data is verified by actual test, but in any case extreme caution should be used for extrapolations that are 10° C below or 20° C above those values tested.

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7/30/97/ АС 27-1A

(5) Engine Power - Turboshaft Engine.

(i) Background.

(A) The purpose of rotorcraft performance flight testing is to obtain accurate quantitative flight test performance data to provide flight manual information.

(B) Flight tests are designed to investigate the overall performance capabilities of the rotorcraft throughout its operating envelope. This testing furnishes information to be included in the flight manual and provides a means of validating the predicted performance of the rotorcraft with a minimum installed specification engine.

(C) The power used to complete the flight manual performance must be based on power values no greater than that available from the minimum uninstalled specification engine after it is corrected for installation losses. A minimum uninstalled specification engine is one that, on a test stand under conditions specified by the engine manufacturer, will produce the certificated power at specification temperatures and/or speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specitied power at a particular engine temperature or speed rating with higher allowable limits. The limit is the maximum value the installed engine is allowed in order to develop the specification power. Prior to installation of each engine in a rotorcraft, the performance is measured by the engine manufacturer. This is done by making a static test run in a test cell and referring the results to standard day, sea level conditions. The performance parameters obtained are presented as uninstalled engine characteristics on a test log sheet. This is commonly referred to as a "final run sheet." Figure 58-3 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 certitied power without exceeding specification operating values; therefore, a "final run sheet" is created for every engine produced. Additionally, if needed, arrangements can usually be made with the engine manufacturer to obtain a torque system calibration for individual engines. This will further optimize the accuracy of the engines used in the flight test program. The engine manufacturer will also provide predicted uninstalled power available for the various power ratings. This information may be derived from an engine computer "card deck" and from charts and tables in the engine detail installation manual. These data also provide engine performance for the range of altitudes and temperatures approved for the engine and include methods for correcting this performance for installation effects. The parameters contained in a typical "card deck" are plotted for one engine rating in Figure 58-4.

(E) Several power losses may be associated with installing an engine in a rotorcraft. Typical losses are air inlet losses, gear losses, air exhaust losses, and powered accessory losses such as electrical generators. Additional flight Par 58 65

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manual performance considerations are the torque indicating system accuracy and torque needle split. The predicted uninstalled power available engine characteristics cannot be assumed to be the actual power available after the engine is installed in the rotorcraft because this procedure would neglect the installation power losses. It is necessary to know the installation losses in order to determine the flight manual performance. Installation losses are reflected reductions in available power resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet and/or exhaust design. The rotorcraft manufacturer conducts tests to confirm the installed specification engine power available on which published performance is based. The specific methods used vary widely between manufacturers but usually include some combination of ground and flight tests. Figure 58-5 is a typical installed power available chart for one set of conditions.

(F) The installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion may occur and there are changes in airflow routing. The rotorcraft manufacturer may elect to determine the installation losses for different flight conditions to take any airspeed advantages. This is acceptable if, for example, the hover performance is based on the actual power available from an installed minimum specification engine in a hover. Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual power available from an installed minimum specification engine at the published climb airspeed. This will allow the manufacturer to take advantage of, for example, increased inlet efficiency.

(її) Procedures.

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

(B) Following a power increase, engine temperature and/or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM and/or temperature are held constant. This is a characteristic typical of turbine engines due largely to expansion of turbine blades and reduced clearances in the engine. Some engines may show a temperature increase at constant power due to engine or temperature sensing system peculiarities. An engine will usually establish a stabilized relationship of power parameters in approximately 2 or 3 minutes. For this reason, the following procedure should be used when obtaining in-flight engine data.

(1) To determine the takeoff and 2 ½-minute values, first stabilize the engine at a low power setting. After stabilization, rapidly increase the power Par 58 66

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demand to takeoff and/or 2 ½-minute power levels. Record the engine parameters as soon as the specification torque, temperature, or speed is attained. Care must be taken not to exceed a limit. These readings should be obtained approximately 15 seconds after power is initially applied.

(2) To determine the 30-minute and/or maximum continuous power values, approximately 2 to 3 minutes of stabilization time after power is increased is generally used, but up to 5 minutes stabilization time is allowed. The reason for the different procedures is when a pilot requires takeoff or 2 ½-minute power values he is in a critical flight condition and does not have the luxury of waiting for the engine(s) to produce rated power. Stabilization time is allowed for the maximum continuous and 30-minute ratings because these values are not associated with flight conditions for which power is needed immediately.

(C) The in-flight measurements recorded with the engine(s) on the flight test rotorcraft must be corrected downward if the test engine is above minimum specification and corrected upward for a test engine that is below minimum specification. This correction is necessary to verify that a minimum specification engine installed on a production rotorcraft is capable of producing the power values used to compute the flight manual performance without exceeding any engine limit. In addition, if the production rotorcraft's power measurement devices have significant (greater than 3 percent) power error, this error must be accounted for in a conservative manner.

(D) On multiengine rotorcraft, the engine location may result in different installation losses between engines. If this condition exists, multiengine performance should be based on the total power available after considering the different installation losses and with minimum specification engines installed.

One-engine-inoperative performance must be based on the loss of the engine which has the lowest installation losses. Additionally, the power losses due to such items as accessory bleed air, particle separators, engine driven accessories, etc., must be accounted for accordingly.

(E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical No speed to a constant corrected speed at cold temperatures. Others may limit power to a fuel flow value which would be encountered only at certain ambients.

Others may limit by torque limiting devices. Therefore, power available data should be obtained at various ambients to verify that all limiting devices are functioning properly and have not been affected by the installation.

(F) Through use, turbine engine power capabilities decrease with time. This is called engine deterioration. Deterioration is largely a function of the particular engine design, the manner, and the environment in which the engine is operated. There is a need, therefore, to provide a method which can be used in service to periodically determine the level of engine deterioration. A power assurance curve is Par 58 67

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AC 27-1A 7130/97 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 58-6 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 No limits that must not be exceeded for a given torque value. An in-flight power assurance check may be used in addition to the pretakeoff check. The validation of either check must be done by the methodology used to determine the installed minimum specification engine power available. For the in-flight power assurance check there must be full accountability for increased efficiency due to such items as inlet ram recovery, absence of exhaust reingestion, etc. A power assurance check done statically and one conducted in-flight must yield the same torque margin(s). An engine may pass power assurance at low power but still may not be capable of producing the rated power values. This occurs when the curve of corrected power and corrected temperature for the engine intersects the minimum uninstalled specification engine curve. If this condition exists, the entire power assurance and power available information may need to be reestablished.

(6) Deteriorated Engine Power - Turboshaft Engine.

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 deteriorated engine power. These include:

(1) Engine characteristics (§ 27.939). Surge margin, engine response, and air-restart capability might be affected and should be addressed, but flight testing may not be required depending on the individual engine/aircraft installation and fuel scheduling mechanism.

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

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

(4) Systems for surge protection which schedule on Pc such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The effect would normally be negligible unless when installed, the installation losses, combined with reduced Pc because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings.

Procedures.

(A) The need for flight tests to verify predicted power available with deteriorated engines depends on the scope of testing which occurred during initial certification. If the original rotorcraft certification included flight testing as described in Paragraph (5) (engine power-turboshaft engines) herein for validation of power available, the need for a demonstration with deteriorated engines is greatly diminished and perhaps eliminated.

(B) If flight testing to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in Paragraph (5) (engine power-turboshaft engines), except that the data would be corrected downward to a deteriorated engine runline. Efforts should concentrate on obtaining data in areas of the operational envelope where maximum gas producer speed is likely to be attained, or where bleed valves or other devices which schedule on gas producer discharge pressure are likely to function. On many installations maximum gas producer speed will occur with cold temperatures and high altitudes; bleed valves and other devices which schedule on gas producer discharge pressure are most likely to function and reduce power on a hot day at low altitude.

(C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated engine monitoring procedures will be introduced when allowing the use of deteriorated power; for example, an in-flight trend monitoring program with the associated bookkeeping duties may be required. Such an in-flight procedure must be evaluated by flight tests as described in Paragraph (5) (engine power-turboshaft engines) herein. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure Par 58 69

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already in place for the specification engine, which could eliminate the need for flight test evaluation.

58A. § 27.45 (Amendment 27-21) GENERAL.

a. Explanation. Amendment 27-21 adds § 27.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered rotorcraft. This information is to provide the pilot a means of determining, prior to takeoff, that each engine will produce the power necessary to achieve the performance presented in the rotorcraft flight manual (RFM).

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

58B. § 27.45 (Amendment 27-29) GENERAL.

a. Explanation. Amendment 27-29 added the requirements for certification of 30-second/2-minute One Engine Inoperative (OEI) power ratings. For rotorcraft approved for the use of 30-second/2-minute OEI, partial power checks currently accomplished with approved power assurance procedures for lower power levels may not be sufficient to guarantee the ability to achieve the 30-second power level.

b. Procedures. Information provided in Appendix 1 of this AC includes guidance material on power assurance procedures to assure that the OEI power level can be achieved.

Par 58 70

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CERTIFICATION BASIS Rgmts 27-Amdt. 21 27-Amdt. 2 FAR 27 Original 6-Amdt. 7 6-Amdt. 4 CAR 6 Original Original CAR 06 7/30 / 97

1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. No specific H-V Level Level Level Level Level Level wt. and alt.

00 Ref. 2. Max. OGE 2. Wt. selected 2. Wt. selected 2. Wt. selected 2. Wt. selected 2. Wt. selected requirements.

27.25 wt. Lesser of: by applicant at by applicant at by applicant by applicant by applicant at

27.1519 27.79 Test b. 7000' Hd a. Max. alt. cap. lessor of: a. Max. alt. cap. lessor of: a. Max. alt. cap. a. Max. alt. cap. at lessor of: Idg. alt. max. t.o. and Idg. alt. max. t.o. and

27.1587 Conditions b. 7000' Hd b. 7000' Hd b. 7000' Hd

6.116

6.741

6.743

3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is H-V is Info. info. info. info. info. operating operating

4. Max. alt. For 4. If H-V wt is limitation. limitation.

REM which H-V is valid less than IGE Wt., wt. becomes limitation for

t.o.Adg.

AC 27-1 5. If H-V is less Para than OGE wt., 58 & 59 H-V wt. becomes limit.

6. Applicant is encouraged to demo H-V to Remarks WAT limits.

7. Hover data may be shown above 7000' if AC 27-1A H-V & IGE are demo'd to 7000' FIGURE 58-1. H-V REQUIREMENTS

### Page 59 (OCR)

CERTIFICATION BASIS FAR 27 CAR 06 Ramts 27-Amdt. 21 Original Original Original

1. MGW Sea Level 1. MGW Sea Level 1. MGW Sea Level 1. No specific CONTROL 27.1587 27.143 27.25 IGE Ref. Conditions Test b. 7000' Hd 2. Max. IGE wt. Lesser of: a. Max. alt. cap. 3. Critical CG 4. Critical Rotor t.o. and Idg. alt. 2. Wt. selected by 3. Critical CG applicant to max. 4. Critical Rotor RPM 2. Max. approved wt. for t.o./dg at alt 3. Critical CG 4. Critical Rotor RPM 5. Wind not less than above sea level. * get ayed in requirement. requirement.

AC 27-1 RPM 5. Critical wt. 20 mph.

Para 5. Wind of not less 6. Wind of not less 58 & 81 than 17 kts. than 20 mph REM 6. Max. safe wind is 7. Max. safe wind 6. Max. safe wind is perf. info. is perf. info. perf. info.

7. If 17 kts. wind demo'd to alt. less Remarks than 7000', a corresponding WAT limit must be established.

**7/30/97 FIGURE 58-2. IGE CONTROLLABILITY REQUIREMENTS**

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**600 CORRECTED SHAFT HORSEPOWER 500 ENG S/N 12345- 400 300 MIN CERTIFIED**

SPECIFICATION ENGINE 200 100 1000 1100 1200 1300 1400 1500 CORRECTED TOT °F FIGURE 58-3. SHAFT HORSEPOWER VS TURBINE OUTLET TEMPERATURE -

SEA LEVEL STANDARD DAY Par 58 73

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**PRESSURE ALTITUDE - FEET (X 1000) 2 200 300 400 500**

SHAFT HORSEPOWER AVAILABLE FIGURE 58-4. UNINSTALLED TAKEOFF POWER AVAILABLE 74 Par 58

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PRESSURE ALTITUDE - FEET (× 1000) 7/30/97 Par 58 FIGURE 58-5. INSTALLED TAKEOFF POWER AVAILABLE, ANTI-ICE OFF, 400 RPM 200 50 ° C 40 ° C SHAFT HORSEPOWER AVAILABLE 300 20 ° C 30 ° C 10 ° C 0°C + -10° € -20 ° C 400 -30 ° 10-50 °C TRANSMISSION AC 27-1A LIMIT 75

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CONDITIONS: EXAMPLE: AC 27-1A ZERO AIRSPEED OAT = 17.5°C RPM 100% TOT = 760° C ANTI-ICE OFF Hp = 4000 FT GENERATOR LOAD 100 AMPS REQUIRED TORQUE = 70% 1000 FT 800°C 8000 FT 780°C 6000 FT 760°C 4000 FT 740°C 2000 FT 720°C SEA LEVEL 700°C 680°C 660°C Par 58 OUTSIDE AIR TEMPERATURE - °C 10 20 30 40 60 20 PERCENT TORQUE 40 60 80 100 120 7/30/97 FIGURE 58-6. POWER ASSURANCE CHECK CHART

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59. § 27.51 TAKEOFF.

a. Explanation. Section 27.51 details the conditions under which takeoff data must be obtained. The flight manual must contain the technique(s) to be used to obtain the published flight manual takeoff procedures. Technique should not be confused with exceptional pilot skill and/or alertness as mentioned in § 27.51. Because rotorcraft differ, different pilot techniques are sometimes required to achieve the safest and most optimum takeoff performance. The recommended technique that is published in the flight manual must be determined to be one that the operational pilot can duplicate using the minimum amount of type design cockpit instrumentation and the minimum crew. Only rotorcraft takeoff techniques will be covered in this section.

b. Background.

(1) Certain special takeoff techniques are necessary when a rotorcraft is unable to take off vertically because of altitude, weight, power effects, or operational limitations. The recommended technique used to take off under such conditions is to accelerate the rotorcraft in-ground-effect (IGE) to a predetermined airspeed prior to climbout. Takeoff tests are performed to determine the best repeatable technique(s) for a particular rotorcraft over the range of weight and altitude for which certification is requested.

(2) Utilizing the total power available to execute a takeoff may not be operationally feasible due to such items as HV or aircraft attitude constraints. In such situations, hover power required plus some power increment may be the maximum recommended for use.

(3) Wheel or skid height should be not less than that demonstrated satisfactorily for the high speed, low altitude portion of the HV curve, or that height below which ground contact may occur when accomplishing takeoff procedures.

(4) For rotorcraft fitted with wheels, a running takeoff procedure may be accepted.

c. Procedures.

(1) There are different takeoff profiles which may be used to complete a maximum performance takeoff in a rotorcraft. The manufacturer will normally determine which method is best for a particular rotorcraft. The most commonly accepted method is the hover and level acceleration technique. In this technique, the rotorcraft is stabilized in a hover at the reference height. From the stabilized hover, the rotorcraft is accelerated to the climbout airspeed using the predetermined takeoff power. When the desired climbout airspeed is achieved, the rotorcraft is rotated and the climbout is accomplished at the scheduled airspeed(s) and constant rotor RPM. Power adjustments may be accomplished to maintain the targeted power except where Par 59 77

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procedure requires high workload outside the cockpit (i.e., that portion of takeoff where horizontal acceleration close to the ground has pilot scan outside the cockpit and adjustment of engine torque or temperature would require an undue increase in workload). The recommended takeoff procedure must be demonstrated to remain clear of the HV "avoid" areas without requiring exceptional piloting skill or exceptionally favorable conditions.

(2) The hover reference height is established as the minimum skid or wheel height above the takeoff surface from which a takeoff can consistently be accomplished in zero wind without contacting the runway surface. The takeoff must be accomplished with power fixed at the power required to hover at the hover reference height and must not require exceptional piloting skill to avoid runway surface contact.

Par 59 78

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**60.-62. RESERVED.**

Intentionally Left Blank Par60 79 (thru 81)

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63. § 27.65 (Amendment 27-14) CLIMB: ALL ENGINES OPERATING.

a. Explanation.

(1) Rotorcraft other than helicopters.

Section 27.65 requires that the steady rate of climb be determined for each rotorcraft other than helicopter with maximum continuous power on each engine for the range of weights, altitudes, and temperatures for which certification is requested.

Equivalent levels of safety have been found wherein the applicant was allowed to select a climb airspeed that was not the actual Vy. The selected airspeed must be consistent with the speed used to show compliance with such items as cooling, stability, etc. The rate of climb resulting from the selected climb airspeed versus that from the actual Vy shall not differ to an extent that a pilot will be encouraged, by appreciable increases in climb performance, to fly a climb airspeed different from that published in the flight manual.

(•) For rotorcraft other than helicopters, the climb performance data obtained above must be used to show that a minimum climb gradient can be achieved for each weight, altitude, and temperature within the range for which certification is required. This gradient must be at least 1:10 if testing is done to determine the required takeoff distance over a 50-foot obstacle. If this option is selected, an explanation of the takeoff distance determination requirements and procedures may be found in Paragraph 62 of AC 29-2.

(iii) If takeoff distance is not determined, the minimum climb gradient must be 1:6 for standard sea level conditions.

(2) For helicopters, Vy must be determined for standard sea level conditions at maximum weight using maximum continuous power on each engine. Although not required, the steady rate of climb may be determined using the procedure in Paragraph 63c.

(3) For helicopters, if VNE at any altitude is less than the maximum gross weight sea level standard day condition Vy, the steady rate of climb must be determined at the climb speed(s) selected by the applicant not to exceed VNE. The climb performance must be determined from 2,000 feet below the altitude from where VNE intersects Vy up to the maximum altitude for which certification is requested. This should be done utilizing maximum continuous power on each engine with the landing gear retracted.

b. Procedure to Determine Vy.

(1) Sawtooth climbs may be used to determine Vy. If such a technique is used, climbs should be flown in pairs on opposite headings 90° to the wind at the test altitude.

This procedure will minimize any windshear effects. All testing must be done in smooth Par63 82

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7/30/97 АС 27-1A air. Windshear is usually an indication of unstable air or a temperature inversion and must also be avoided. The climbs are flown on reciprocal headings for approximately 5 minutes or through an altitude band using maximum continuous power at a constant airspeed. Periodic power adjustments may be necessary. Additional reciprocal heading climbs must also be conducted at different airspeeds above and below the airspeed at the lowest point of the power required versus airspeed curve. This technique can be repeated at different altitudes to obtain Vy throughout the altitude range.

(2) Level flight performance (speed power) may also be used to determine Vy.

The testing should be done in smooth air. The advantage of this method is that less time is required, and the accuracy is equivalent to the sawtooth climb method. The test can be repeated at various altitudes to determine the Vy throughout the altitude range desired for the rotorcraft. The test at each altitude should be conducted at a constant weight over sigma (W/o). The test is normally started at the desired W/o with maximum continuous power, or at VNE, in level flight. A series of points should be taken, reducing airspeed 10 to 15 knots between points, with the lowest speed point around 20 to 30 knots. Weight should be computed for each point and the test altitude adjusted to maintain a constant W/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 may be reached prior to a limiting torque. The test team should verify that the resulting power utilized in these tests closely approximates the power producing capabilities of a minimum installed specification engine.

c. Procedure to Determine All-Engine-Operating Climb Performance.

(1) Background. Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to validate the rotorcraft's climb performance.

By-products are a qualitative evaluation of the rotorcraft handling characteristics in a climb and engine data to assist in the determination of installed power available.

(2) Techniques. The climbs are conducted on reciprocal headings at the established airspeed(s) through the target altitude range. The same parameters are recorded as during sawtooth climbs. The rotorcraft will usually climb very rapidly during the first few thousand feet; therefore, the data acquisition method must be timely if accurate results are expected. This procedure is usually repeated at weight extremes.

The resulting data must then be corrected for power and weight. Power and weight Par 63 83

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corrections are satisfactory, provided the test powers and weights closely approximate the target values to make the weight and power corrections small. Once this data is finalized and corrected for all the flight test variables, interpolation for intermediate weights can be made with a high degree of reliability. If the rotorcraft has any stability augmentation system, vent systems, etc., which may influence the climb performance, then it must be accounted for. Caution should be taken that anti-ice, air-conditioning, etc., are not on unless the performance is being established specifically for those conditions.

63A. § 27.65 (Amendment 27-33) CLIMB: ALL ENGINES OPERATING.

a. Explanation. Amendment 27-33 added the requirement to determine the steady rate of climb, for helicopters, from sea level up to the maximum altitude for which certification is requested. Although not specifically stated in the rule, the rate of climb should be determined at Vy or, if VNE at any altitude is less than the maximum gross weight sea level standard day condition Vy, the steady rate of climb at these altitudes must be determined at a climb speed(s) selected by the applicant not to exceed VNE-

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

64. § 27.67 (Amendment 27-23) CLIMB: ONE ENGINE INOPERATIVE.

a. Explanation.

(1) Section 27.67 requires that for multiengine normal category rotorcraft, the steady rate of climb or descent with one engine inoperative must be determined at Vy (or at the speed for minimum rate of descent) for maximum gross weight.

(2) The rate of climb (or descent) will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or 30-minute minimum specification installed power available values. The landing gear should be retracted if it is retractable.

b. Procedures

(1) The procedure discussed in Paragraph 63 for all-engines-operating climb performance is also applicable to the OEI condition. For twin-engine rotorcraft that are shown not to have a "critical engine" with respect to performance characteristics, both engines may be used to simulate the appropriate single-engine power available during these tests.

(2) Adequate testing must be accomplished to determine the rotorcraft's OEI climb performance at maximum gross weight for all variations in altitude and temperature for inclusion in the Rotorcraft Flight Manual.

84 Par 63

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65. § 27.71 (Amendment 27-21) GLIDE PERFORMANCE.

a. Explanation.

(1) Performance capabilities during stabilized autorotative descent are useful tools to assist the pilot when all engines fail. This information is also useful in determining the suitability of available landing areas along a given route segment.

(2) Two speeds are of particular importance, the speed for minimum rate of descent and the speed for best angle of glide. These speeds along with glide distance information are required as flight manual entries per § 27.1587. The speed for minimum rate of descent is useful for engine failure conditions at higher altitudes and the pilot is required to perform some time-related task, engine restart, float inflation, radio calls, etc. The speed for best angle of glide is a somewhat higher speed that is of particular use when it is necessary to reach a distant landing area. These speeds, when utilized in conjunction with appropriate rotor RPM and glide angle (or rate of descent) can be used to calculate the maximum horizontal distance available from a particular altitude assuming zero wind conditions.

(3) A third speed, recommended autorotation speed, may be provided in addition to minimum rate of descent speed and maximum glide angle speed. The recommended speed for autorotation is usually optimized to assure an effective flare capability and yet be slow enough to allow a controlled, relatively slow touchdown condition. Recommended autorotation speed is ordinarily between the minimum rate of descent and maximum glide angle speeds. The recommended autorotation speed may be provided in the Rotorcraft Flight Manual. The relationship between minimum rate of descent, best glide angle, and recommended autorotation speed is shown in Figure 65-1.

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

b. Procedures.

(1) Tests are conducted at speeds which bracket the anticipated speeds for minimum rate of descent and best glide angle. On a power required plot, the speed for minimum power required approximates the speed for minimum rate of descent. The speed for maximum range glide may be estimated by drawing a tangent from the origin to the power required curve.

(2) Autorotative performance tests may be conducted in conjunction with the climb performance tests. The required data are similar for both tests and it is sometimes convenient and efficient to run alternating climbs and descents through a Par 65 85

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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) During autorotation tests, care must be taken to make certain that no engine power is delivered to the rotor drive system since a very small amount of power can have a large effect on descent performance.

Par 65 86

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**RATE OF**

**DESCENT RECOMMENDED**

**AUTOROTATIONAL SPEED**

**MINIMUM BEST GLIDE ANGLE**

**RATE OF DESCENT**

AIRSPEED FIGURE 65-1. AUTOROTATIONAL CHARACTERISTICS - TYPICAL Par 65 87

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66. § 27.73 PERFORMANCE AT MINIMUM OPERATING SPEED.

a. Explanation.

(1) The word "hover" applies to a rotorcraft that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air. For the purpose of this manual, the word "hover" will mean pure hover.

(2) The regulatory requirement for hover performance, § 27.73, refers to hover in ground effect (IGE). For some applications, such as external load operations, hover performance out-of-ground effect (OGE) is necessary; however, it is not required by this section. Hover OGE is that condition, where an increase in height above the ground will not require additional power to hover. Hover OGE is the absence of measurable ground effect. It can be less than one rotor diameter at low gross weight increasing significantly at high gross weight. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle one and one-half rotor diameters above the surface.

(3) The objective of hover performance tests is to determine the power required to hover at different gross weights, ambient temperatures, and pressure altitudes.

Using nondimensional power coefficients (Cp) and thrust coefficients (Cy) for normalizing and presenting test results minimizes the amount of data required to cover the rotorcraft's operating envelope.

(4) Hover performance tests must be conducted over a sufficient range of pressure altitudes and weights to cover the approved ranges of those variables for takeoff and landing. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.

(5) The hover ceiling for which data should be obtained and subsequently presented in the flight manual should be the same height consistent with the minimum hover height demonstrated during the takeoff tests. Refer to Paragraph 59 for the procedure to determine this hover height.

b. Procedures.

(1) Two methods of acquiring hover performance data are the tethered and the free flight techniques. The tethered technique is accomplished by tethering the rotorcraft to the ground using a cable and load cell. The load cell and cable are attached to the ground tie-down and to the rotorcraft cargo hook. The load cell is used to measure the rotorcraft's pull on the cable. Hover heights are based on skid or wheel height above the ground. During tethered hover tests, the rotorcraft should be at light gross weight. The rotorcraft will be stabilized at a fixed power setting and rotor speed at the appropriate skid or wheel height. Once the required data are obtained, power should be varied from the minimum to the maximum allowed at various rotor RPM. This 88 Par 66

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technique will produce a large C/C spread. The load cell reading is recorded for each stabilized point. The total thrust the rotor produces is equal to the rotorcraft's gross weight plus the weight of the cables and load cell plus cable tension. Care must be taken that the cable tension does not exceed the cargo hook limit or load capacity of the tie-down. For some rotorcraft, it may be necessary to ballast the rotorcraft to a heavy weight in order to record high power hover data.

(2) The pilot maintains the rotorcraft in position so that the cables and load cell are perpendicular to the ground. To ensure the cable is vertical, two outside observers, one forward of the rotorcraft and one to one side, can be used. Either hand signals or radio can be used to direct the pilot. The observers should be provided with protective equipment. Positioning can also be accomplished by attaching two accelerometers to the load cell which sense angle or movement along the longitudinal and lateral axes.

Any displacement of the load cell will be reflected on instrumentation in the cockpit, and by reference to this instrumentation, the rotorcraft can be maintained in the correct position. Increased caution should be utilized as tethered hover heights are decreased because the rotorcraft may become more difficult to control precisely. The tethered hover technique is especially useful for OGE hover performance data because the rotorcraft's internal weight is low and the cable and load cell can be jettisoned in the event of an engine failure or other emergency.

(3) To obtain consistent data, the wind velocity should be less than 3 knots as there are no accurate methods of correcting hover data for wind effects. Rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are:

(i) Engine torque.

(il) Rotor speed.

(III) Ambient temperatures.

(iv) Pressure altitude.

(v) Fuel used (or remaining).

(vi) Load cel reading.

(vii) Generator(s) load.

(vili) Wind speed and direction.

As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover data should be verified by a limited spotcheck using the free flight technique. The free Par 66 89

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flight technique as contained in Paragraph 66b(4) will determine if any problems, such as load cell malfunctions, have occurred. The free flight hover data must fall within the allowable scatter of the tethered data.

(4) If there are no provisions or equipment to conduct tethered hover tests, the free flight technique is also a valid method. The disadvantage of this technique as the primary source of data acquisition is that it is very time consuming. In addition a certain element of safety is lost OGE in the event of an emergency. The rotorcraft must be reballasted to different weights to allow the maximum Cy Cp spread. When using the free flight technique, either as a primary data source or to substantiate the tethered technique, the same considerations for wind, recorded parameters, etc., as used in the tethered technique apply. Free flight hover tests should be conducted at CG extremes to verify any CG effects. If the rotorcraft has any stability augmentation system which may influence hover performance, it must be accounted for.

(5) It is extremely difficult to determine when a rotorcraft is hovering OGE at high altitudes above ground level since there is no ground reference. In a true hover, the rotorcraft will drift with the wind. Numerous techniques have been tried to allow OGE hover data acquisition at high altitudes, all of which have resulted in much data scatter. Until a method is proposed and found acceptable to the FAA/AUTHORITY, OGE hover data must be obtained at the various altitude sites where IGE hover data are obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet.

67. § 27.75 (Amendment 27-14) LANDING.

a. Explanation.

(1) This rule incorporates all of the landing requirements for Part 27 rotorcraft.

(2) As with other flight maneuvers, landings must be accomplished with acceptable flight and ground characteristics using normal pilot skills. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in Paragraph 58 of this advisory circular. As in other performance areas, engines must be operated within approved limits.

(3) Landing. Approach and landing path requirements are stated in general terms in Paragraphs (a)(1) and (a)(2) of § 27.75. The approach path must allow smooth transition for a one-engine-inoperative landing and adequate clearance from potentially hazardous HV combinations.

(4) All-engine-out landing. Section 27.75(b) contains the certification requirement for "last" engine failure and all-engines-inoperative landing. The rule states that it must be possible to make a safe landing after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously, although Par 66 90

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complete power failure has occurred in twin-engine rotorcraft with Category A engine isolation. This requirement assures that in the event of cockpit mismanagement, fuel exhaustion, improper maintenance, fuel contamination, or unforeseen mechanical failures, a safe autorotation entry can be made and a safe power-off landing can be affected. Two separate aspects of this rule are normally evaluated at different times during the test program. The "last" engine failure is normally evaluated during cruise or VNE engine failure testing where instrumentation and critical loading have been established for those test conditions. The all-engine-out landing is ordinarily conducted in conjunction with an HV or landing distance phase where ground instrumentation and safety equipment are available.

b. Procedures.

(1) Instrumentation/Equipment. Aircraft instrumentation may include engine and flight parameters, control positions, power lever position, and landing gear loads. A record of rotor RPM at touchdown is necessary to assure it does not exceed transient limits. Rotor RPM at touchdown may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. A crash recovery team with the support of a fire engine is highly desirable.

(2) The one-engine-inoperative landing is similar in many respects to the HV tests described in paragraph 69 of this advisory circular. Most of the comments, cautions, and techniques for HV also apply here even though the typical flight conditions are less critical than limiting HV points due to a lower power level and an established rate of descent. The approach is made at a predetermined speed with one engine inoperative. The speed is reduced and the rotorcraft is flared to a conventional one-engine-inoperative landing.

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

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

(4) Aircraft Loading. Aft center of gravity is usually most critical because visibility constraints limit the degree to which the pilot can see the landing surface during the flare. If a weight effect is shown, a minimum of two weights should be flown at each test altitude. One weight should be the maximum weight for prevailing conditions, and the other should provide a sufficient spread to validate weight accountability.

(5) All-engine-out landing.

Par 67 91

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(i) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion of this paragraph, § 27.75(b) contains two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. The second aspect of this rule requires that a landing from autorotation be possible. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. If a complete company test program has documented an all-engine-out landing to the GWI (gross weight/density ratio) limit, verification tests may be initiated at those limiting weight conditions. If not, buildup testing should be initiated at light weight. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings-one at sea level and one near maximum takeoff and landing altitude.

(її) Demonstrated compliance with this requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to a reasonable value (less than 40 KTAS is recommended) to ensure a reasonable chance of survivability for the all engine failure condition. On multiengine rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to this relatively low inertia level, considerable collective may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which may lead to rotorcraft damage in testing this 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.

(ії) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train.

Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished.

Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a multiengine rotorcraft totally without power. The all-engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.

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7/30/97 АС 27-1A

(6) Prior to conducting these tests, the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons idle power be used if a "needle split" (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure needle split is attained. In other cases a temporary detent between idle and cutoff was used on the throttle. In a third case the engine was actually shut down on sample runs to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as is reasonable to dissipate speed and build RPM Rotor RPM must stay within allowable limits. Aft center of gravity is ordinarily critical due to visibility and flarability. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it must be within allowable structural limits.

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7/30/97 AC 27-1A

**68. RESERVED.**

Intentionally Left Blank Par 68 94

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69. § 27.79 (Amendment 27-14) LIMITING HEIGHT-SPEED ENVELOPE.

a. Explanation.

(1) The height-speed envelope is normally referred to as the height-velocity (HV) diagram. It defines an envelope of airspeed and height above the ground from which a safe power-off or OE 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 69-1. 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 Vu) and make a successful landing or climb out on the remaining engine(s).

(2) Power failure, engine failure, throttle chop, or other similar terms used in this discussion mean a simulated engine failure. The actual shutdown of an engine to simulate an engine failure should not be necessary if the simulated procedure ensures that the engine power is suddenly removed from driving the rotor and remains so. The normal fuel control deceleration schedule is usually satisfactory for the power removal for turbine engines but the flight/ground idle speed may have to be set lower than normal for HV testing.

(3) The avoid areas of the HV diagram are separated by the takeoff corridor.

This corridor should be wide enough to consistently permit a takeoff flight path clear of the HV diagram using normal pilot skill. The takeoff corridor should always permit a minimum of +5 knots clearance from critical portions of the diagram.

(4) The knee of the curve separates the takeoff portion from the cruise portion and is defined as the highest speed point on the low speed portion of the HV envelope.

Altitudes above this point are considered cruise, or "fly-in," points, and these test points require a minimum time delay of 1 second between throttle chop and control actuation (reference § 27.143(d)). Altitudes below the knee represent takeoff profile points. For test points in the takeoff portion, takeoff power (or a lower power selected by the applicant as an operating procedure) and normal pilot reaction time for corrective control actuation will be used.

(5) Since the HV diagram may represent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer's buildup program should be reviewed to determine the amount of conservatism in the HV diagram (if any). It should be remembered that the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points 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, Par 69 95

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AC 27-1A 7130/97 the ability to arrest and control descent rates through use of an aft cyclic flare may be greatly diminished. Extreme care should be exercised when "backing down" to lower speeds.

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

(7) For FAA/AUTHORITY tests, the minimum required crew and the minimum instrument panel display presented for certification should be used. Ground safety equipment should be provided.

(8) This test is the least predictable of all the performance items. Therefore, the expansion and extrapolation of test data are questionable. Weight may not be extrapolated to higher values. In order to extrapolate HV data to higher altitudes, any analytical method must have FAA/AUTHORITY approval. In lieu of pure analytical methods, simulations have been used successfully, especially for multiengine rotorcraft.

In either case, the maximum allowable extrapolation should be limited to 2,000 feet density altitude (Ho). HV test weights for normal category rotorcraft are the maximum weight at sea level and some lessor weight at high density altitudes. The high density altitude HV curve needs to be defined only to 7,000 feet and may be a lower altitude if the rotorcraft does not have the performance capabilities to attain 7,000 feet. A weight less than the maximum weight may be used to define the high density altitude HV curve, but this weight should not be less than the maximum weight that will allow hovering out-of-ground effect. For a given diagram, typical weight reductions that are necessary as altitude is increased can be conservatively estimated by maintaining a constant gross weight divided by density ratio, GW/o. See Figure 69-2, 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 69-2, Part B. Another method of presentation is to show varying weights at a constant density altitude. (See Figure 69-2, 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, landing, and HV tests.

b. Procedures.

(1) Instrumentation.

Ground Station. The ground station must have equipment and instrumentation to determine wind direction and velocity, outside air temperature, and if the test rotorcraft has reciprocating engines, humidity. Since the tests must be conducted in winds of 2 knots or less, a smoke generator is highly recommended to show both flightcrew and ground crew personnel the wind direction and velocity at any given time. Additionally, the location of the ground station should be such that it is free Par 69 96

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of rotor downwash at all times. Motion picture or phototheodolite and radio equipment will be necessary to properly conduct the test program. The use of telemetry equipment is desirable if the location of the test site and the magnitude of the test program make it practical.

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

(2) Analytical Prediction. The HV diagram can be estimated by analytical means and this is recommended prior to test. HV, however, is the least predictable of all rotorcraft performance and because of this, the expansion and extrapolation of test data must be done with great care. Test weight may not be extrapolated. All test points should be approached conservatively with some speed or altitude margin. If the applicant has conducted a comprehensive HV flight test program to validate his analytical predictions, much preliminary testing can be eliminated. In any case, the maximum allowable extrapolation from flight test conditions is 2,000 feet density altitude, and an approved analytical and/or simulation method must be utilized for extrapolation.

(3) Power.

The appropriate power level before engine failure for the low and high hover points is simply the power required to hover at the prevailing hover conditions.

The appropriate power condition prior to failure of the engine for points below the knee is takeoff power or a lower value if approved as an operating procedure. For cruise or "fly-in" points above the knee, the appropriate condition is power required for level flight.

(ii) The applicable power failure conditions are listed in § 29.79(b).

Power should be completely cut for normal category rotorcraft. For multiengine rotorcraft with Category A engine isolation, only one engine need be failed and the desired topping power (for the remaining engine(s)) should be set prior to the test. This power value will need adjustment as ambient conditions change. The power can be takeoff power (TOP), 2 ½-minute power, or some calculated lower power for simulating hot day or higher density altitude conditions. Power is verified and recorded by the pilot by "topping" the engine(s) prior to engine failure tests. Care must be taken to ensure that this power value is no more than that which would be delivered by a minimum specification engine under the ambient conditions to be approved.

(4) Test Loadings. Weight extrapolation is not permitted for HV. Therefore, the test weight must be closely controlled. Ballast or fuel should be added frequently to Par 69 97

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maintain the weight within -1 to +5 percent when testing final points. Ordinarily, tests are conducted at a mid center of gravity unless a particular loading is expected to be particularly critical.

(5) Landing Gear Loads.

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.

(її) Any landing which results in permanent deformation of aircraft structure or landing gear beyond allowable maintenance limits is considered an unsatisfactory test point.

(6) Piloting Considerations. In verifying the HV diagram, the minimum certificated instrument panel display and minimum crew should be used in order not to mislead the operational pilot who has no test equipment available and may have no copilot to assist. Three distinctly different flight profiles are utilized in developing the diagram.

(i> High Hover. A stabilized out-of-ground-effect (OGE) hover condition prior to power failure is essential. A minimum 1-second time delay between power failure and initial control actuation is utilized. Following the time delay, the primary concern is to quickly lower collective and to gain sufficient airspeed to allow an effective flare approaching touchdown. While the immediate development of airspeed is necessary, the dive angle must be reasonable and must be representative of that expected in service. While initial aircraft attitude will vary between models and with changing conditions, 10°-20° has been previously applied as a maximum allowable nose down pitch attitude. Use of greater attitudes could result in a diagram which is difficult to achieve and unrealistic for operations in service. Initial testing should start relatively high with gradual lowering of height to the final high hover altitude. A stabilized OGE hover condition prior to power failure is essential. If a stabilized high hover condition cannot be achieved prior to the engine cut, then this point should be tested from a minimum level flight speed. This will result in an open-ended HV diagram. A smoke source or balloon on a long cord is highly desirable since the wind can vary significantly from surface observations to typical high hover altitudes. Vertical speed must be very near zero at the throttle chop. Any climb or sink rate can have a significant influence on the success of the test point. Use of a radar altimeter with a cross check to barometric altitude is essential.

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 Par 69 98

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7/30/97 АС 27-1A lowering of collective immediately after power failure can result in a very high, unconservative low hover height that is unrealistic for operational conditions. If, however, a design is such that a 1-second pilot delay after power failure could be achieved without any appreciable descent, a slight lowering of collective could be allowed.

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

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

(7) Ground Support. Motion picture or theodolite coverage and ground safety equipment are necessary. Communication capability among these elements should be provided. Use of a phototheodolite to compare height/speed with cockpit observations is very desirable.

(8) Verifying the HV Diagram.

A sufficient number of test points must be flown to verify the diagram.

The key areas are the knee, high altitude hover, low altitude hover, and low altitude high speed flight. Test points with excessive gear loads, exceptional skill requirements, winds above permissible levels, rotor droop below approved minimum transient RPM, damage to the rotorcraft, excessive power, incorrect time delay, etc., cannot be accepted.

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

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

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(10) Night Evaluation. If a rotorcraft is to be certified for night operation, a night evaluation is required. Simulated engine failures should be conducted along the recommended takeoff path. Landings should also be qualitatively evaluated with an engine failed. Engine failures at critical HV conditions are not required. The intent is to show adequate visibility using aircraft and/or runway lights without requiring a duplication of the daytime HV test program.

(11) Water Landings. For amphibious float-equipped rotorcraft, day and night water landings should be conducted under critical loading conditions with an engine failed. Engine failures should be conducted along the recommended takeoff path.

Engine failures at critical HV conditions are not required. The intent is to show similarity to test results over land without requiring a duplication of the HV test program.

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**7/30/97/ AC 27-1A HEIGHT LEVEL FLIGHT (H)**

- KNEE

**TAKEOFF**

**TAKEOFF CORRIDOR**

**HIGH SPEED**

SPEED (V) FIGURE 69-1. HEIGHT - VELOCITY DIAGRAM Par 69 101

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• Approx.

Constant GWo Height Velocity LA- Weight

CONSTANT HV DIAGRAM, VARIABLE WEIGHT Increasing Density Altitude Height Velocity

- B

CONSTANT WEIGHT Increasing Gross Weight Height Velocity

- C-

CONSTANT DENSITY ALTITUDE FIGURE 69-2. ALTITUDEMEIGHT ACCOUNTABILITY Par 69 102

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69A. § 27.79 (Amendment 27-21) LIMITING HEIGHT-SPEED ENVELOPE.

a. Explanation. Amendment 27-21 to the regulation redefines the required weight for establishing the HV envelope at altitudes above sea level.

b. Procedures. All of the policy material pertaining to this section remains in effect. In addition, the following applies:

(1) The rotorcraft height-velocity envelope should be established for the maximum gross weight at sea level. At altitudes above sea level, the envelope should be established at not less than the maximum operating weight or OGE hover weight, whichever is lower. If a weight below the OGE hover weight is selected, by definition, that selected weight becomes the maximum operating weight for the rotorcraft at that altitude.

(2) If the HV envelope is established for a maximum altitude less than 7,000 feet, by definition, the maximum takeoff and landing altitude for the rotorcraft may be no higher than that maximum HV altitude. Hover performance information should not be presented for altitudes above the maximum altitude for which the HV envelope is established.

**70.-79. RESERVED.**

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**SECTION 4. FLIGHT CHARACTERISTICS**

80. § 27.141 (Amendment 27-21) GENERAL.

a. Explanation.

(1) This section prescribes the general flight characteristics required for certification of a normal category rotorcraft. Specifically, it states that the rotorcraft shall comply with the flight characteristics requirements at all approved operating altitudes, gross weights, center of gravity locations, airspeeds, power, and rotor speed conditions for which certification is requested. The reference to altitude in § 27.141(a)(1) refers to density altitude. Density altitude is, of course, a function of pressure altitude and ambient temperature, hence the need to account for ambient temperature effects.

Additional flight characteristics required for instrument flight are contained in 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 §§ 27.141 through 27.161 are met at takeoff power levels, no flight manual entry is needed. Obviously if transmission limits for maximum continuous (MC) and takeoff power coincide, no special action is needed.

(3) During the flight characteristics testing, the controls must be rigged in accordance with the approved rigging instructions and tolerances. The control system rigging must be known prior to testing. In addition to the normal rigging procedures, any programmed control surfaces which may be operated by dynamic pressure, electronics, etc., must also be calibrated. During the flight test program, it is frequently necessary to rig a control, such as the swashplate or tail rotor blade angle, to the allowable critical extreme of the tolerance band. For example, it would be necessary to rig the tail rotor to the minimum allowable blade angle if meeting the requirements of §27.143(c) would be in question. The same consideration must be given to all rotorcraft controls and movable aerodynamic surfaces where questionable compliance with the regulations may exist. If the rotor-induced vibration characteristics of the rotorcraft are significantly affected and require time-consuming rigging for such things as acceptable ride comfort, then the rotor(s) should be rigged to the allowable extreme tolerance limits to determine compliance, for example, with § 27.251.

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7/30/97 АС 27-1A

(4) During the FAA/AUTHORITY flight test program, the crew should be especially alert for conditions requiring great attentiveness, high skill levels, or exceptional strength. If any of these features appear marginal, it is advisable to obtain another pilot's opinion and to carefully document the results of these evaluations.

Section 27.141(b) provides the regulatory basis for these strength and skill requirements. The general requirements for a smooth transition capability between appropriate flight conditions are also included in § 27.141(b). These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft.

(5) For night or IFR approval, § 27.141 (c) contains the general regulatory reference which requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this order.

81. § 27.143 (Amendment 27-21) CONTROLLABILITY AND MANEUVERABILITY.

a. Explanation.

(1) This regulation contains the basic controllability requirements for normal category rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for control and for maneuverability are summarized in § 27.143(a) which is largely self-explanatory. The hover condition is not specifically addressed in § 27.143(a)(2) so that the general requirement may remain applicable to all rotorcraft types, including those without hover capability. For rotorcraft, the hover condition clearly applies under "any maneuver appropriate to the type."

(2) Paragraphs (b) through (e), § 27.143, include more specitic flight conditions and highlight the typical areas of concern during a flight test program.

(i) Section 27.143(b) specifies flight at VNE With critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at VNE for nosedown pitching of the rotorcraft and for adequate roll control. Nosedown pitching capability is needed for control of gust response and to allow necessary flight path changes in a nosedown direction. Roll control is needed for gust response and for normal maneuvering of the aircraft. In the past, 10 percent control travel margin has been applied as an appropriate minimum control standard. The required amount of control power, however, has very little to do with any fixed percentage of remaining control travel. There are foreseeable designs for which 5 percent remaining is adequate and others for which 20 percent may not be enough. The key is, can the remaining longitudinal control travel at VNE generate a clearly positive nosedown pitching moment, and will the remaining lateral travel allow at least 30° banked turns at reasonable roll rates? Moderate lateral control reversals should be included in this evaluation and since available roll control can diminish with sideslip, reasonable out of trim conditions (directionally) should be investigated. This "control remaining" philosophy must also be applied for other flight conditions specified in this section.

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(it) Section 27.143(c) requires a minimum control capability for hover and takeoff in winds of 17 knots from any azimuth. Control capability in wind from zero to at least 17 knots must also be shown for any other appropriate maneuver near the ground such as rolling takeoffs for wheeled rotorcraft. These requirements must be met from standard sea level conditions to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less. On rotorcraft incorporating a tail rotor, efficiency of the tail rotor decreases with altitude so that a given sideward flight condition requires more pedal deflection, a higher tail rotor blade angle, and more horsepower. Hence, directional capability in sideward flight (or at critical wind azimuth) is most critical during testing at a high altitude site.

Section 27.143(d) requires adequate controllability when an engine fails. This requirement specifies conditions under which engine failure testing must be conducted and includes minimum required delay times.

(A) For rotorcraft which meet the engine isolation requirements of transport Category A, demonstration of sudden complete single-engine failure is required at critical conditions throughout the flight envelope including hover, takeoff, climb at Vy, and high speed flight up to VNE. Entry conditions for the first engine failure are engine or transmission limiting maximum continuous power (or takeoff power where appropriate) including reasonable engine torque splits. For multiengine Category A installations (three or more engines) subsequent engine failures should be conducted utilizing the same criteria as that used for first-engine failure. The applicant may limit his flight envelope for subsequent failures. Initial or sequential engine failure tests are ordinarily much less severe than the "last" engine failure test required by § 27.75(b).

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

(B) For rotorcraft without transport Category A engine isolation, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to VNE (power-on) and conditions of hover, takeoff, and climb at Vy. Maximum continuous power is specified prior to the failure for the cruise condition. Power levels appropriate to the maneuver should be used for other conditions. The corrective action time delay for the cruise failure should be 1-second or normal pilot reaction time (whichever is greater). Cyclic and directional control motions which are apart of the pilot task of flight path control are normally not subject to the 1-second restriction; however, the delay is always applied to the collective control for the cruise failure. If the aircraft flying qualities and cyclic trim configuration would encourage routine release of the cyclic control to complete other cockpit tasks during cruise flight, consideration should be given to also holding cyclic fixed for the 1-second delay. Although the same philosophy could be extended to the directional controls, the likelihood of the pilot having his feet away from the pedals is 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 116 Par 81

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7130/97 AC 27-1A "cruise" also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies maximum continuous (MC) power, it does not limit engine failure testing to MC power. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions. Following power failure, rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.

(iv) Section 27.143(e) addresses the special case in which a VNE (power-off)is established at an airspeed value less than VNE (power-on). For this case, engine failure tests are still required at speeds up to and including VNE (power-on), and the rotorcraft must be capable of being slowed to NE (power-off) in a controlled manner with normal pilot reactions and skill. There is, however, no controllability requirement for stabilized power-off flight at speeds above 1.1 VNE (power-off) when VNE (power-off) is established per § 27.1505(c).

(v) Application of the controllability requirement for pitch, roll, and yaw at speeds of 1.1 VNE (power-off) and below is similar to that described above for power-on testing at VNE. Sufficient directional control must exist to allow straight flight in autorotation during all approved maneuvers including 30° banked turns up to VNE (power-off) with some small additional allowance for gust control. Adequate controllability margins must exist in all axes throughout the approved autorotative flight envelope. Testing to VNE at MC power per § 27.143(b), 1.1 VNE at power for 0.9 VH per § 27.175(b) or § 27.1505, and to 1.1 VNE (power-off) in autorotation per § 27.143(e) should be sufficient to assure adequate control margin during a descent condition at high speed and low power. The high speed, power-on descent condition should be checked for adequate control margin as a "maneuver appropriate to the type." There has been one instance where insufficient directional pedal was available to maintain a reasonable trimmed sideslip angle with low power at very high speeds, and a case where there was insufficient forward and lateral cyclic available to reach the power on VNE. The insufficient directional pedal margin was due to the offset vertical stabilizers.

The lack of cyclic stick margin was because the cyclic stick migrated to the right as power was reduced, and the control limits were circular. This provided less total available forward cyclic stick travel when the cyclic was moved right and forward about 45° from the center position. Each of the above rotorcraft was certificated with a rate of descent limitation to preclude operation in the control-limited area.

(vi) An evaluation of the emergency descent capability of the rotorcraft should be made, either analytically or through flight test. Areas of consideration are the rate of descent available, the maximum approved altitude, and the time before a catastrophic failure following the loss of transmission oil pressure or other similar failure. Each rotorcraft should have the capability to descend to sea level and land from the maximum certificated altitude within the time period established as safe following a critical failure. If the time period does not permit a sea level landing, the maximum height above the terrain must be specified in the limitation section of the Rotorcraft Flight Manual.

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(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 27.672). Examples include stability augmentation systems (SAS), stability and control augmentation systems (SCAS), automatic flight control systems (AFCS), devices to provide or improve longitudinal static stability such as a pitch bias actuator (PBA), yaw dampers, and fly-by-wire elevator or stabilator surfaces. These systems all use actuators of some type, and they are subject to actuator softover and hardover malfunctions. The flight control system should be evaluated to determine whether an actuator jammed in an extreme position would result in reduced control margins.

Generally, if the flight control system stops are between the actuator and the cockpit control, the control margin will be affected. If the control stops are between the actuator and the rotor head, the control margins may not be affected, but the location of the cockpit control may be shifted. This could produce interference with other items in the cockpit. An example of this would be a lateral actuator jammed hardover causing a leftward shift in the cyclic stick position. Interference between the cyclic stick, the pilot's leg, and the collective pitch control could reduce the left lateral control available and reduce left sideward flight capability. In the case of fly-by-wire surtaces, 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 VNE controllability testing is ordinarily aft CG, MC power, and minimum power-on rotor RPM, although power and RPM variations should be specifically evaluated to verify their effects. The turbine engine is sensitive to ambient temperatures which affect the engine's ability to produce rated maximum continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit maximum continuous power would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward CG should be spot checked for any "tuck under" tendency at high speed. The VNE controllability test is normally accomplished shortly after the 1.1 VNE (or 1.1VH) point obtained during stability tests 118 Par 81

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required by § 27.175(b). Controllability must be satisfactory for both conditions. If VNE varies with altitude or temperature, VNE for existing ambient conditions is utilized for the test. Extremes of the altitude/temperature envelope should be analyzed and investigated by flight test.

(3) The critical condition for controllability testing in a hover is ordinarily forward CG at maximum weight with minimum power-on rotor RPM. For rearward flight testing of configurations where the forward CG limit varies with weight, low or high gross weight may be critical. Lateral CG limits should also be investigated. A calibrated pace vehicle is needed to assure stabilized flight conditions. Surface winds should be less than 3 knots throughout the test sequence. Testing can be done in higher stabilized wind conditions (gusting less than 3 knots); however, these conditions are very difficult to find and the method is very time consuming due to the necessity of waiting for stabilized winds. Testing in calm winds is preferred. Hover controllability testing should be accomplished with the lowest portion of the rotorcraft at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. Although the necessary yaw response will vary somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the rotorcraft is held in the most critical position relative to wind.

(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 § 27.903(b). The propulsion engineer should have assured that this feature was properly addressed in the engine and drive system substantiation;

however, it must be assumed that for some period of time the pilot may extract maximum available power from the remaining engine(s) when an engine fails during critical flight maneuvers. Substantiation of this feature should be accomplished primarily by engine and drive system ground tests.

(5) Longitudinal cyclic authority at VNE With any power setting must permit suitable nosedown pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft's response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary Par 81 119

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recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft motion, the results are unacceptable; e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral cyclic to correct roll in conjunction with forward cyclic to correct pitchup is satisfactory.

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

(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all-engine-inoperative requirements of § 27.75(b).

See Paragraph 67 of this AC.

82. § 27.151 (Amendment 27-21) FLIGHT CONTROLS.

a. Explanation. Excessive breakout or preload in the flight controls produces control system force discontinuities which result in increased workload and 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. In some designs friction can provide a positive contribution to the function of the flight controls (e.g., masking aerodynamic teedback in reversible systems). At some point, friction will have a detrimental effect on the pilot's ability to properly control the machine. In the case of an irreversible design equipped with an artificial force feel system in pitch and roll, excessive friction can mask a shallow force gradient making positive stick centering and control force static stability difficult if not impossible to demonstrate. In such an instance, the initial choice of fixes might include implementation of a steeper force gradient or addition of a force preload. Unfortunately, these solutions often lead to the kinds of problems discussed earlier. Care must therefore be exercised during the initial design phase to ensure that the components and characteristics of the flight control system are well matched.

b. Procedures. Regardless of the flight control system sophistication, it is important that the test pilot understand the system configuration prior to flight evaluation. Appropriate mechanical characteristics should be documented. For VFR aircraft, the mechanical characteristics are typically assessed in flight on a qualitative basis. If a controllability or workload problem is identified, a more detailed investigation would be necessary. Since IF 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.

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

83. § 27.161 (Amendment 27-21) TRIM CONTROL.

a. Explanation.

(1) The pilot has many tasks to perform with each hand during sustained flight conditions. The trim requirement is intended to reduce the physical demands to maintain a given flight condition. It is not intended to require that control forces be reduced to zero by the trim control during dynamic maneuvers such as takeoff acceleration.

(2) A number of devices may be used to produce the necessary trim characteristics. One popular method of meeting this requirement is through the use of control balance springs in conjunction with a small amount of built-in control system friction. Other methods include use of friction, magnetic brakes, bungees, and irreversible mechanical schemes.

(3) This regulation is not intended to require zero friction or zero breakout force in the control system, nor is it intended to require automatic control recentering. The regulation, in fact, specifically prohibits excessive high friction or high breakout forces which would produce undesirable discontinuities in the primary control force gradient.

b. Procedures.

(1) If comprehensive company flight test data are available, compliance with this requirement can quickly be found by spot checking extreme center of gravity loadings. Trim tests can ordinarily be done during the course of other flight test activities. To conduct the test, briefly release the control at the required flight conditions and determine that the control does not move. The words "any appropriate speed" ordinarily include any speed from hover to VH. If the control system trim device might be subject to temperature or humidity effects, these should be investigated at a minimum of two altitude extremes and during several test phases.

(2) If a pilot controllable variable friction device is incorporated, compliance with this requirement must be shown at the minimum adjustable value. The maximum value of adjustable friction should not completely lock the flight controls.

(3) Continued compliance with this requirement should be ensured through a production procedure. If minimum friction or centering springs are used, it is desirable for the manufacturer to include some adjustment capability for production differences.

The explanation and procedures discussed here are applicable for VFR approval under § 27.161. For additional FR trim requirements, refer to Paragraph 775 of this advisory circular.

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АС 27-1A 7/30/97

84. § 27.171 STABILITY: GENERAL.

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

b. Procedures.

(1) Compliance with the requirements of this section can often be obtained for the VFR condition without any specific or designated flight testing. If the rotorcraft is marginal in regard to pilot strain and fatigue, the FAA/AUTHORITY pilot should be assured, through special tests if necessary, that the aircraft can be satisfactorily flown throughout the maximum endurance capabilities of the rotorcraft including night and turbulence conditions if those are critical. This test should be conducted with minimum required systems in the aircraft and with minimum flightcrew.

(2) Reasonable failure conditions which add to pilot workload, strain, and fatigue should be evaluated (electrical, hydraulic, and mechanical failures, etc.). The necessary times associated with flight with a failed system must be appropriate to the flight manual procedures for each failure. A failure condition requiring immediate landing would obviously require shorter evaluation time than a condition allowing continued flight to destination.

(3) IFR approvals necessitate a careful evaluation of Paragraphs b (1) and (2) above. In FR operations, weather conditions frequently necessitate continued flight to destination or diversion to alternate airports with critical failures. Immediate landing may not be feasible. The evaluating pilot must ensure pilot strain and fatigue are acceptable during typical flight profiles for each type of operation to be approved.

85. § 27.173 (Amendment 27-21) STATIC LONGITUDINAL STABILITY.

a. Explanation.

(1) This rule contains control requirements for both stability and control.

Paragraph (a) contains the basic control philosophy necessary for all civil aircraft.

Forward motion of the cyclic control must produce increasing speeds, and aft motion must result in decreasing speeds. For rotorcraft, this is accomplished with throttle and collective held constant. Rotorcraft with either highly stable or highly unstable static longitudinal stability characteristics can typically comply with the basic requirement for control sense of motion. However, the intent and interpretation of this paragraph is to provide a stable stick position versus airspeed gradient. Therefore, a stabilized airspeed less than the trim speed requires a cyclic stick position aft of the trim stick position, and a stabilized airspeed greater than the trim speed requires a cyclic stick position forward of the trim speed stick position.

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(2) The remainder of § 27.173, through reference to § 27.175, contains the basic control position requirements necessary to establish a minimum level of static longitudinal stability. Positive stability is found for conditions of climb, cruise, and autorotation in § 27.175 by requiring a stable stick position gradient through a specified speed range. A defined level of instability is permitted for the hovering condition.

b. Procedures.

(1) The control requirement of this section is so essential to basic flight mechanics that compliance may be found during conventional flight testing for compliance with other portions of the regulations. No special or designated testing should be required.

(2) The procedures necessary to ensure compliance with the stability requirements of this section are contained under § 27.175. Refer to Paragraph 86 of this advisory circular for an explanation of detailed flight test procedures.

86. § 27.175 (Amendment 27-21) DEMONSTRATION OF STATIC LONGITUDINAL STABILITY.

a. Explanation.

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

(2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in § 27.173 which states that throttle and collective pitch must be held constant is appropriate for administration of this rule, and the specified power in § 27.175(a), (b), and (c) should be considered as power established at initial trim conditions. This will result in slightly higher or lower torque readings at "off trim" conditions. Collective and throttle controls are held constant when obtaining data during climb, cruise, and autorotation tests.

(3) The effects of rotor RPM on autorotative static stability should be determined and positive stability demonstrated for the most critical RPM. Values for RPM can be expected to change as airspeed is varied from the "trimmed" condition.

The manufacturer's recommended autorotation airspeed is ordinarily used for trim.

(4) Hovering is considered a flight maneuver for which the pilot repeatedly adjusts collective to maintain an approximately constant altitude above the ground. For hover stability tests, collective and throttle adjustments are made as necessary to Par 85 123

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АС 27-1А 7/30/97 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.

Sensitive control position instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when using a "before and after" method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The Figure 86-1 plot is a typical presentation of longitudinal static stability.

(її) Other necessary parameters include pressure altitude, ambient temperature, and indicated airspeed (pace vehicle or theodolite speed for hover tests).

For hover tests, hover height (radar altitude if available) and surface winds should be documented. Two-way communication with a pace vehicle is highly desirable. Ground safety equipment is desirable.

(2) Ambient Conditions. Smooth air is necessary for stability testing. Allowable wind conditions for hover stability testing are the same as those for hover controllability tests. Extrapolation is covered in Paragraph 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.

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

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(i¡) The cruise test 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 flight trimmed condition at 0.9 VH (or 0.9 VNE if lower). This point is then recorded as the trim point. The collective pitch and throttle must remain fixed at the trim setting for the remainder of the test. The airspeed is then varied above and below the trim speed using the cyclic control to climb or dive slightly.

(iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually ‡2,000 feet), first increasing airspeed as data points are collected, then decreasing speed through the same altitude band. It will probably not be possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit. Generally, it will be possible to acquire all the high speed points on one pass and the low speed points on the second.

(iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise, data can be obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise 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 (0.9Vu) trim speed, 125, 145, 115, 155, 105, and 165.

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

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

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

(vii) Normally, climb, cruise, and autorotation tests should be conducted at low, medium, and high altitudes. See Paragraph 58 for guidance on interpolation and extrapolation. High speed stability has been critical during cold weather testing. In two recent models, VNE at cold temperatures has been limited by the stability requirements Par 86 125

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АС 27-1A 7/30/97 of § 27.175(b). Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather VNE to tip Mach number values demonstrated during warm weather testing.

(viii) Hover stability should be verified at low altitude and, if required, at high altitude. Refer to Paragraph 58b(2) for guidance on expansion and extrapolation of altitude.

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**O CRUISE**

• CLIMB

**A AUTOROTATION**

• HOVER

SHADED POINTS REPRESENT TRIM 100 90 0000000 000000 LONGITUDINAL CYCLIC POSITION (% FROM FULL AFT) • 20 10 20 40 60 80 100 120 140 160

CALIBRATED AIRSPEED (KNOTS) FIGURE 86-1. STATIC LONGITUDINAL STABILITY Par 86 127

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87. § 27.177 (Amendment 27-21) STATIC DIRECTIONAL STABILITY.

a. Explanation. This rule requires that positive static directional stability be demonstrated at the trim airspeeds defined in § 27.175. The trim speed for climb is Vy and for cruise 0.9VH or 0.9VNE (whichever is less).

b. Procedures.

(1) Tests for static directional stability require instrumentation for pedal position and sideslip angle. To obtain accurate sideslip angle and airspeed information, a "yaw boom" is usually installed for the purpose of mounting a sideslip vane and swiveling airspeed pilot head outside the main rotor downwash region of influence. Special care should be taken to ensure that the yaw boom installation has been verified to be structurally adequate and free of dynamic instabilities for all combinations of airspeed and rotor speed likely to be experienced during the static directional evaluation. For some installations, the instrumentation yaw boom may influence the flying qualities of the rotorcraft itself. Thus, it is advisable to correlate yaw string displacement or slip indicator ball widths of skid with yaw boom sideslip angle, and then repeat a few critical points with the yaw boom removed.

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

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

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

(Ш1) Smoothly apply directional control and coordinate with lateral control to establish the desired sideslip angle. A steady track can best be ensured by maintaining a track over a straight landmark on the ground such as a section line or straight segment of powerline or highway.

(iv) Because drag increases with sideslip, the aircraft will decelerate. The trim airspeed should be maintained by entering a slight dive or decreased rate of climb.

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If a boom airspeed system is not used, a "standard" airspeed system will become excessively inaccurate at about 10° of sideslip. Rapidly yawing to the desired sideslip angle and noting and maintaining the new "standard" indicated airspeed may give adequate data. Control positions (directional as a minimum), sideslip angle, rotorspeed, airspeed, rate of descent, amount of ball deflection, and bank angle should be recorded. The pilot should note the physical sideforce feel experienced. The rule requires that sufficient cues accompany sideslip to alert the pilot when approaching sideslip limits. A minimum of two sideslip data points on each side of the trim point should be obtained to adequately define the slope of the pedal displacement versus sideslip angle relationship.

(v) Static directional stability 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.

**88.-95. RESERVED.**

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**SECTION 5. GROUND AND WATER HANDLING CHARACTERISTICS**

96. § 27.231 GENERAL.

a. Explanation. The rule states: "The rotorcraft must have satisfactory ground and water handling characteristics, including freedom from uncontrollable tendencies in any condition expected in operation.i" In addition, §§ 27.235, 27.239, and 27.241 contain specific requirements concerning ground and water handling characteristic evaluations.

b. Procedures.

(1) During the flight test program and the F&R program (§ 21.35(b)(2)), the rotorcraft will be subjected to evaluations at various weight and CG conditions. Any uncontrollable tendencies found during these test programs must be corrected.

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

(3) Any significant vibration or 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 speed and at CG extremes. If a nose or tail wheel lock/swivel control is installed, each position should be evaluated for limiting takeoff, landing, and taxi speeds. Maximum substantiated speed values should be included in the RFM as limitations.

(5) For water operations, the wave 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 RFM. Information or limits on the allowable "sea state" for rotor startup and shutdown should also be included.

97. § 27.235 TAXING CONDITION.

a. Explanation. The rotorcraft is designed for certain landing load factors (§§ 27.471 and 27.473). The rotorcraft must not attain a load factor in excess of the design load factor when taxied over the roughest ground that may reasonably be expected in normal operation at the expected taxi speeds. This rule applies to wheel landing gear equipped rotorcraft.

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b. Procedures. The structural substantiation data contain the allowable design limits for the rotorcraft. A calibrated accelerometer or load factor "g" meter should be installed as near as practicable to the rotorcraft CG to record the maximum vertical load factor attained. Instrumentation of the landing gear and/or related structure may also be an acceptable means of showing compliance.

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

(2) The taxi surface should be evaluated for compliance with the rule.

Corrugated surfaces as well as broken or uneven surfaces (in accordance with the rule) should be used.

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

(4) A light and heavy rotorcraft 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. § 27.239 SPRAY CHARACTERISTICS.

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

b. Procedures.

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

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Con- Rotor Altitig. 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, if any.

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 § 27.609 and Paragraph 245 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. § 27.241 GROUND RESONANCE.

a. Explanation.

(1) The rule states: ¿®The rotorcraft may have no dangerous tendency to oscillate on the ground with the rotor turning." This rule is a flight requirement that pertains to demonstrating freedom from dangerous oscillations on the ground. CAR Part 6, predecessor to FAR Part 27, originally contained a "strength requirement" under § 6.203 requiring ground vibration tests. These tests would identify critical vibration frequencies and modes of the rotorcraft. CAR Part 6, Amendment 6-4, effective October 1, 1959, removed this ground vibration requirement because the agency concluded that if any major component has a natural frequency which could be excited by some operating parameter, such a condition would be revealed in the course of other ground and flight tests. The FAA/AUTHORITY apparently was depending on demonstrations under § 6.131/§ 27.241 and the flight load survey data (§ 27.571) to 140 Par 98

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7/30/97 АС 27-1A satisfy the objective of the vibration test. However, Part 27, Amendment 27-2, contained new § 27.663 adding reliability and damping action investigation requirements for ground resonance prevention means. A ground vibration survey was not reinstituted by the adoption of § 27.663. Compliance with § 27.663 does require investigation and substantiation as stated.

(2) "Ground resonance" is a mechanical instability of the aircraft while in contact with the ground, often when partially airborne. Stated another wayiground resonance" is a self-excited mechanical instability that involves coupling between the in-plane motion of the rotor blade and the motion of the rotorcraft as a whole on its landing gear (reference "Aerodynamic of the Helicopter," Gessow & Myers, page 308).

It is caused by the motion of the blade in the plane of rotation (called in-plane vibration) coupled with a rocking or vertical motion of the aircraft as a whole. The tires, landing gear, and rotor pylon restraint structure act as a spring with a vibration frequency which coincides or couples with the natural in-plane frequency of the blade about a real or effective drag hinge in the plane of rotation. When the frequencies of the two motions (rotor and airframe) approach each other and couple, a violent shaking of the rotorcraft may occur which, if undamped, could result in the destruction of the rotorcraft.

(3) Ground resonance can occur due to 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.

(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 two-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.

(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 warmup. Their damping characteristics require thorough investigation for the range of rotorcraft operating environment as noted in § 27.663.

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b. Procedures.

(1) Under all conditions, any oscillations which may be introduced should be damped. However, no instability should occur at any operating condition such as during RPM changes from minimum to maximum and idle to maximum. For rotorcraft with wheel gear, uneven taxi surfaces in conjunction with particular taxi speeds, may excite ground resonance and should be evaluated by taxiing on typical surfaces. This evaluation may be conducted in conjunction with the tests of § 27.235. In operation, the resonance characteristics should be checked during takeoff and landing at zero speed and during run-on landings using various power values.

(2) For those aircraft equipped with 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 and 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. The collective pitch setting increments should range from flat pitch to light on the landing gear prior to fully airborne, depending upon the test sequence for minimum risk.

(5) Cyclic pitch inputs should be made 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 obtain representative responses, typically time sufficient to execute five complete circles of the cyclic stick (about neutral) at the selected frequency.

(6) The excitation frequency should be such as to excite the blade in-plane frequency. Rotor speed settings should be increased to 1.05 times the maximum power-on rotor speed. Collective pitch settings should be increased in increments of 142 Par 99

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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) × 300 RPM) or 3.5 cycles per second to attempt excitation of possible resonant frequencies. At the conclusion of the excitation, the cyclic stick should be returned to the neutral position while continuing the recording of data listed in Paragraph b(13).

(8) The excitation process should include cyclic excitation inputs from the directional and longitudinal controls if critical for the type of rotorcraft being evaluated.

(9) If onset of ground resonance is encountered, one possible corrective action is to increase the collective pitch and rotor speed and become airborne. However, lowering the collective pitch and applying the rotor brake (if installed) or rolling off the throttles has been effective for some designs and is considered a satisfactory procedure if resonance can be consistently stopped.

(10) With the rotor speed stabilized, landing should be made at a touchdown speed which minimizes risk.

(11) Special Considerations.

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

Additionally, the potential of misservicing and possible failure modes should be evaluated. For ground resonance qualification, where practical, variations from the baseline test configuration may be accomplished by ground run (§ 27.663(b) requires investigation of probable ranges of damping), analyses, component tests, aircraft shake test, the specification of special operational procedures in the rotorcraft flight manual, or a combination thereof. Detailed and rational analyses showing acceptable correlation to the baseline tests, and for which the input parameters were verified by drawings, calculations, component static or dynamic tests, or by aircraft shake tests simulating the conditions/configurations in question, may be used to limit testing to only those variables and operational conditions showing marginal or unacceptable system damping. All operational limitations should be clearly stated in the rotorcraft flight manual. A report of the analytical results and/or test results should be submitted per § 27.663.

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

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

(13) Instrumentation and Data Acquisition.

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

(А) Altitude.

(В) OAT.

(C) Wind velocity.

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

(А) Gross weight.

(B) C.G.

(C) Tire pressure.

(D) Landing gear oleo pressure.

Instrumentation (for recording during test).

(А) Main rotor RPM.

(B) Time history of cyclic control fore-and-aft and lateral stick position.

(C) Time history of collective control stick position.

(D) Time history of rotor damper motion.* (Е) Time history of pylon component motion.* (F) Time history of landing gear (oleo) motion.* (G) Time history of aircraft motions.* *As required to obtain modal damping 144 Par 99

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**100.-108. RESERVED.**

Intentionally Left Blank Par 145 (thru 153)

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**109. RESERVED.**

Intentionally Left Blank 154 Par 109

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**SECTION 6. MISCELLANEOUS FLIGHT REQUIREMENTS**

110. § 27.251 VIBRATION.

a. Explanation.

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

(2) This flight requirement may be both a qualitative and quantitative flight evaluation. Section 27.571(a) contains the flight load survey requirement that results in accumulation of vibration quantitative data. Section 27.629 generally requires quantitative data to show freedom from flutter for each part of the rotorcraft including control or stabilizing surfaces and rotors.

(3) Review Case No. 70 (reference FAA Order 8110.6) contains a policy statement concerning compliance with this rule. This policy statement is condensed here for convenience:

"The rotorcraft must be capable of attaining a 30° bank angle (turn), at VNE, with maximum continuous power (maximum continuous torque) without encountering excessive roughness/vibration. The FAA/AUTHORITY requires the maneuver demonstration to provide the pilot with some maneuver capability at VNE and further to provide the pilot some margin away from roughness when operating in turbulencei (This maneuver may result in a descent or a climb.)

(4) Section 27.1505 pertains to VNE determination. Section 27.1509 pertains to rotor speed limits determination.

b. Procedures.

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

The flight loads survey quantitative data (reference § 27.571) and the applicant's qualitative and quantitative flight test data must also prove compliance with the requirement prior to issuing an authorization for official FAA/AUTHORITY flight tests.

(2) The flight load survey data obtained under § 27.571(a) will contain measured data concerning proof of freedom from flutter and excessive vibration.

Pertinent critical flight conditions will be reinvestigated during FAA/AUTHORITY flight tests. The specific condition or conditions necessary to demonstrate compliance with § 27.251 vary with the rotorcraft design and with the minimum and maximum rotor Par 110 155

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

(3) The airspeed and rotor speed limits investigated and established under §§ 27.33, 27.1503, 27.1505, and 27.1509 are also investigated and made a matter of record in the flight loads survey data. During the official FAA/AUTHORITY/TIA flight tests, critical parts of the rotorcraft may have limited instrumentation to reinvestigate and confirm that the critical conditions investigated during the flight load survey are satisfactory and do not result in excessive vibration. Use of instrumentation is optional if the flight loads data are conclusive.

(4) FAA/AUTHORITY policy for 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/AUTHORITY requires maneuver demonstrations of 30° banked turns at NE Without encountering excessive roughness. The maneuver should be conducted with the rotor speed at the minimum RPM and maximum RPM limits.

During the flight load survey, this condition should be investigated and data recorded to ensure hazardous loads are not encountered for this "unusual" condition. As indicated, the flight condition will be reinvestigated during the FAA/AUTHORITY flight tests. See Paragraph 110b(2) for illustration of this speed and RPM demonstration.

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• Bank, Autorotation 105% Max RPM Power-Off 4 4 Level Forward Flight, Power-On

• 30° Bank, Power-On Max RPM Power-Off Max RPM Power-On 46 ROTOR RPM Min RPM Power-On д5 Min RPM Power-Off 07 95% Min RPM Power-Off 03 VNE 1.11 VNE

**AIRSPEED**

1. Autorotation at 1.11 VNE (AR) at minimum placard rotor speed.

2. Autorotation at 1.11 VNE (AR) at maximum placard rotor speed.

3. Autorotation at VNE (AR) at power-off minimum design limit rotor speed.

4. Autorotation at VNE (AR) at power-off maximum design limit rotor speed.

5. Forward flight 1.11 VNE at minimum power-on rotor speed.

6. Forward flight 1.11 VNE at maximum power-on rotor speed.

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

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

Note: VNE (AR) may be less than VNE • FIGURE 110-1. DEMONSTRATION POINTS Par 110 157

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**111.-119. RESERVED.**

Intentionally Left Blank 158 (thru 166) Par 111

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AC 27-1A 7/30/97

120. RESERVED Intentionally Left Blank 167 Par 120

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**SECTION 7. STRENGTH REQUIREMENTS - GENERAL**

121. §27.301 LOADS.

a. Explanation.

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

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

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

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

122. § 27.303 FACTOR OF SAFETY.

a. Explanation.

(1) Unless otherwise provided by Part 27, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will ensure that the design strength of the rotorcraft is greater than the design loads contained in Part 27.

(2) Other rules, §§ 27.561(b)(3) and 27.787(c), specify use of defined ultimate inertial forces for protection of occupants.

b. Procedures.

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

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

123. § 27.305 STRENGTH AND DEFORMATION.

a. Explanation.

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

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(2) The structure is required to be able to support, in a static test, ultimate loads for 3 seconds without failure, or dynamic tests simulating actual load application may be used.

(3) Section 27.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by structural analysis. See Paragraph 124 of this AC.

b. Procedures. Any test results, static or dynamic, should 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 Part 27 should contain the criteria stated in the rule.

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

124. § 27.307 (Amendment 27-3) PROOF OF STRUCTURE.

a. Explanation.

(1) The rule requires compliance with the strength and deformation requirements for each critical loading condition. Certain tests must be conducted as specified. Additional tests for new or unusual design features may be required as noted in § 27.307(b)(6).

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

b. Procedures.

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

(2) Whenever tests are used or required, a test proposal or plan should be approved prior to the tests. The test article should have received conformity inspections and should have been accepted by the FAA/AUTHORITY for the test. Test fixtures and instrumentation should also be acceptable to the FAA/AUTHORITY (using DERs as appropriate) prior to the start of the test. The quality control office of the applicant or other qualified personnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FAA/AUTHORITY or DER performing Par 123 169

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АС 27-1A 7/30/97 this task. The test proposal may be used to define and to authorize the means to accomplish inspection of the test fixtures and instrumentation. Unnecessary drawings such as test fixture details or layering of approvals are not intended or envisaged by this policy. Drawings, sketches, or photographs have been used by the FAA/AUTHORITY to control and to ensure correct location, direction, and magnitude of loads and other critical test parameters.

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

Generally, the rotorcraft airframe should have natural frequencies remote to predominant rotor excitation sources, including higher harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this vibration. During the flight load measurement program conducted under § 27.571, critical loaded areas or critical joints may be instrumented with strain gages or other stress strain measuring devices. This actual flight data should be compared to the analytical data to verify accuracy.

(4) Paragraph (b) of the rule specifies certain tests. Test proposals should be approved prior to conducting official FAA/AUTHORITY tests. Other paragraphs in this advisory circular pertain to those tests.

124A. § 27.307 (Amendment 27-26) PROOF OF STRUCTURE.

a. Explanation. Amendment 27-26 adds the requirement to account for the environment to which the structure will be exposed in operation. This change is intended to codify recent FAA/AUTHORITY and industry practices for the consideration of environmental effects in showing "proof of structure."

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

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

(2) MIL-HDBK-5, AC 20-107B, 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. § 27.309 DESIGN LIMITATIONS.

a. Explanation.

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

(2) Refer to the rule 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 necessary to revise structural design criteria limits to agree with more conservative operational limits established during the certification program. The operational limits may be subsequently expanded by additional flight tests to agree with design limits.

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**126.-134. RESERVED.**

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**135. RESERVED.**

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**SECTION 8. FLIGHT LOADS**

136. § 27.321 (Amendment 27-11) GENERAL.

a. Explanation.

(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 gravityi

b. Procedures.

(1) Derivation of the flight loads is required by and specified in §§ 27.337 through 27.351. This rule requires flight load determination from minimum to maximum weight and for disposable loads.

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

137. § 27.337 (Amendment 27-26) LIMIT MANEUVERING LOAD FACTOR.

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

(1) A range of design positive load factors from +3.5 to +2.0 may be used.

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

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

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

b. Procedures.

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(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.

The maximum positive design load factor of +3.5 is generally at a weight below maximum gross weight. The maximum thrust capability of the main rotor, combined with incremental lift of wings or sponsons, if installed, results in a maximum design positive load factor. An example of a load factor-gross weight curve is shown in Figure 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 § 27.321(b)(1) since the design load factor varies with gross weight.

(її) 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 § 27.337(b), the flight demonstration required by § 27.337 (b) must be conducted. The flight test personnel should determine that the demonstration shows the probability of exceeding the selected design load factors (those factors less than +3.5 and more than -1.0) is extremely remote. (See Order 8110.4, Paragraph 166c(2)(c)).

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

Par 137 183

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+3.5 Represents curve of Design thrust + lift Load Factor

• +2.0 Nz -1.0 Minimum Gross Weight Lbs. Maximum

LOAD FACTOR GROSS WEIGHT CURVE FIGURE 137-1 184 Par 137

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138. § 27.339 (Amendment 27-11) RESULTANT LIMIT MANEUVERING LOADS.

a. Explanation. The rule specifies or defines the application of rotor and lift surtace loads to the rotorcraft.

(1) The design maneuvering load factors required by § 27.337 will result in or be derived from rotor thrust or lift and from auxiliary surface lift.

(2) Sections 27.321, 27.337, and 27.341 all complement one another and result in the derivation of design flight loads that will be imposed to ensure structural integrity of the rotorcraft.

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

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

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

(Il) 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 $$ 27.321, 27.337, and 27.341. These loads are applied to the rotor or rotors and any auxiliary surface as prescribed by this rule.

139. §27.341 GUST LOADS.

a. Explanation.

(1) The rotorcraft must be substantiated for the loads derived from 30 feet per second vertical gusts from hovering to 1.11 VNE (i.e., VD).

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AC 27-1A 7130/97

(2) Gust loads for any horizontal stabilizing surface should be derived for vertical gusts, upward and downward.

b. Procedures.

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

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) 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 is generally negligible when compared to the rotor and may be neglected for the rotorcraft gust load determination if proven negligible by analysis.

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

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

(V) Gust loads on the stabilizing surfaces are required as stated in Paragraph 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 may be attributed to the small incremental change in lift due to the 30 FPS gust. Nonetheless, design gust loads for the rotorcraft shall be calculated as specified in the rule to ensure the rotorcraft maneuvering load factors do, in each case, exceed the design gust load factor.

(3) For further information about rotorcraft gust response characteristics, see Paper No. 9 presented at the AHS/NASA -Ames Specialist's Meeting on Rotorcraft Dynamics, February 13-15, 1974. The paper, entitled, "Helicopter Gust Response Characteristics Including Unsteady Aerodynamics Stall Effects,i was written by P.J.

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

140. § 27.351 (Amendment 27-26) YAWING CONDITIONS.

a. Explanation. The standard was added by Amendment 27-26. It requires proof of a rotorcraft "structural" yaw or sideslip design envelope. This sideslip envelope 186 Par 139

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should cover minimum forward speed or hover to the lesser of VH or VNE for "power-on" condition, not "power-off" since Vy is a part of the standard. The rotorcraft should be structurally safe for the thrust capability of the directional control system as stated.

(1) The rotorcraft structure should 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 derived from § 27.397(a)), is required for the conditions cited in the standard. A control system rate limiter or a yaw damper may be used as part of the type design required equipment, if elected. The total displacement is therefore a function of time as well as the maximum effort applied such as 130 pounds.

(i> At low airspeeds from 0 to 0.6 VNE, 90° yaw (sideward flight) is specified as the design limit.

(Ii) At high airspeeds VH or VNE), stabilized yaw angle (stabilized sideslip) or 15° sideslip, whichever is less, is specified to be substantiated.

(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 § 27.351(b)(1).

(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.

b. Procedures.

(1) Many of the current single main rotor rotorcraft designs have vertical (directional) stabilizing surfaces. These surfaces may be solely vertical stabilizing fins as on the Bell Model 206 or a swept vertical extension of the tail boom as on the Hiller Model FH1100. The Hiller FH1100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.

(1) 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 FH1100, the tail surface air loads will add to or subtract from the tail rotor thrust according to the flight condition under consideration.

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NOTE: For one example: At stabilized yaw to the right (left pedal depressed to limit) (§ 27.351(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 § 27.351(b)(3), the tail rotor thrust moment then is added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 27.351(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 § 27.351(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 logical assessment of the various yaw conditions may be used to reduce the load derivation and analysis workload for critical rotorcraft design conditions.

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

(vii) A simple structural design envelope may be derived from these design data. If the right or left yaw limits are not very different, common and 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 0.6 VNE) and 15° yaw when they are applied at the lesser of VNE or VH, with a straight line variation from 0.6 VNE to VH or VNE. The rotorcraft does not need to be capable of attaining the 90° and 15° yaw. They should be considered as the maximum sideslip limits.

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90°

YAW DEGREES (LONGITUDINAL AXIS) → 15° MAX.

0°

**0.6 VNE AIRSPEED**

**LESSER OF**

VA OR VNE FIGURE 140-1. SAMPLE YAW/FORWARD SPEED DIAGRAM Par 140 189

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(vili) During flight test evaluations, yaw angles have been measured using a yaw angle probe (swiveling vane type) on a nose boom. Both a 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. However, this test is not a requirement.

(2) FAR § 27.351(b)(1) incorrectly references § 27.395(a) for the maximum pilot forces. The correct reference should be § 27.397(a).

141. § 27.361 (Amendment 27-23) ENGINE TORQUE.

a. Explanation.

(1) The rotorcraft shall be designed for limit engine torque values, as prescribed by the rule, to account for maximum engine torque, including certain transients and torsional oscillations. Amendment 27-23 separated the standard into paragraphs for turbine and reciprocating engine limit torque values.

(2) Turbine engine limit torque for design purposes (Amendment 27-23) was redefined into four cases and the torque values determined will be used. For example, sudden engine stoppage is introduced as one of the cases which is applied to the engine and the engine suspension and restraint system. Emergency operation of governor-controlled turboshaft engines is another case.

(3) Torque factors are also specified for reciprocating engines having two or more cylinders in Paragraph (b) of the standard.

(4) Sections 27.547(e)(1)(ii) and 27.549(d), respectively, refer to the application of engine torque to design of main rotor structure and engine mount and adjacent structure.

b. Procedures.

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

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

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(3) Turbine engine limit torque values are determined for the four cases specified. Two cases are related to the endurance test of §§ 27.923 and 27.927.

(4) For sudden stoppage of turbine engines the engine manufacturers can reasonably provide FAA/AUTHORITY approved data to the applicant on inertia of rotating parts and the deceleration time expected in the event of sudden engine stoppage. This condition usually generates critical loads in the engine mounting and restraint system. These manufacturer's data should be acceptable for use in compliance with this part of the standard.

142.-151. RESERVED Par 141 191 (thru 201)

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**SECTION 9, CONTROL SURFACE AND SYSTEM LOADS**

152. § 27.391 GENERAL.

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

b. Procedures. The design criteria and/or the design loads report shall contain the loads dictated by the referenced rules. (See Paragraphs 153, 154, 155, 156, 157, 158, and 159 of this document.)

152A. § 27.391 (Amendment 27-26) GENERAL.

a. Explanation. Amendment 27-26 adds an explicit reference to § 27.427, Unsymmetrical Loads (AC Paragraph 160), to clarify that substantiation for unsymmetrical loads is a general control surface requirement. A reference to § 27.399, Dual Control System (AC Paragraph 155), is also added for clarification. In addition, §§ 27.401, 27.403, and 27.413 were removed by this amendment since these references and requirements were adequately addressed in other standards.

b. Procedures. The referenced AC paragraphs become 153, 154, 155, 158, and

160.

153. § 27.395 CONTROL SYSTEM.

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

(1) Paragraph (a) of the rule specifies the way or means of reacting the minimum design loads specified in §§ 27.397 and 27.399 (for dual control systems).

Except reduced design loads, not less than 0.60 of those specified in §§ 27.397 and

27.399 for dual control system, may be used as specified. The standard also applies to those control systems that may have more than one stop in a system. The design loads must be imposed on the system from the pilot's control to any stop in the control system.

(2) Minimum design loads imposed on the control system from a stop to a rotor blade or a control surface or device shall be:

(i) The maximum pilot forces obtainable in normal operation; and If low operational loads may be exceeded as noted in § 27.395(b)(2), the system shall support without yielding 0.60 of the loads specified in §§ 27.397 and

27.399 for dual control systems.

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(3) Section 27.695 concerns standards for a power boost and power-operated control system. This standard, in effect, imposes a fail-safe standard for hydraulic aspects of a control system. Where appropriate to a particular design, the control system must therefore sustain without yielding, the maximum output force of the actuator when complying with § 27.395(a). The pilot input forces are not added to the actuator output forces according to this standard for normal category rotorcraft. These forces are independently applied to the control system.

(4) Control system design features and tests requirements are found in §§ 27.619 and 27.625, respectively. Special factors such as casting, bearing, and fitting factors that may be appropriate for the design are contained in §§ 27.619 and

27.625, respectively.

b. Procedures.

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

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

A static test proposal for testing the control system to show compliance with the rules should be approved before conducting the test. Where compliance is to be determined by tests, limit load tests, as discussed in Paragraph 284 of this document, and/or ultimate load tests may be performed. Test results shall be documented.

(ii) If tests are not conducted, a structural analysis of the control system is required. Appropriate factors from §§ 27.623 and 27.625 must be used as specified.

Tests may not be required when adequate similarity of systems and support structure is determined and where adequate structural analysis is furnished.

(3) If a part of the control system is not stiff or rigid enough to react the design loads specified in §§ 27.397 and 27.399, that part of the system may be substantiated for lower loads as prescribed.

(i) The limit design loads are those loads specified in §§ 27.397 and

27.399;

The maximum that can be obtained in normal operation and that is allowed by the system; except (її) The limit design loads may not be less than 0.60 of the limit pilot forces specified.

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(iv) For example, if a small control surface or servo tab is lightly loaded, its control system must be stiff enough to react the control surface loads and to provide surface deflection to control the rotorcraft. The normal operational loads may be very low, such as 10 pounds maximum. Nonetheless, the design limit load shall be

0.60 times the limit single pilot forces specified in § 27.397. Note that the system must not yield under these loads.

(v) For example, if a dual but primary manual control system such as a tail rotor control is lightly loaded, the control system, from the stops to the rotor blades, may be designed for minimum loads equal to 0.60 times the limit dual pilot forces specified in § 27.399.

(vi) If a power actuator is a part of a rotor control system, the design limit force for the affected parts shall be the maximum output force of the actuator at any operational condition (including any load/pressure after a single failure in the hydraulic system).

(4) Controls proof and operation test is required by §§ 27.307(b)(2) and (b)(3),

27.681, and 27.683. This test is conducted using the design limit loads approved under § 27.395. (See Paragraphs 284 and 285 of this document.)

153A. § 27.395 (Amendment 27-26) CONTROL SYSTEM.

a. Explanation. Amendment 27-26 extensively rewrites § 27.395(b) to more clearly incorporate design condition loads for typical powered control systems. New requirements include substantiation for loads resulting from "each normally energized power device, including any single power boost or power activator failure in the control system. There are also new minimum loads for control system designs in which operational loads may be exceeded through jamming, ground gusts, control inertia, or friction. The old loads were 0.60 times the limit pilot forces of § 27.397; the new loads are 100 percent of the limit pilot forces specified in § 27.397.

b. Procedures. The procedures of Paragraph 153 continue to apply except that the increased loads in § 27.395(b)(4) of 100 percent of limit pilot forces are specified for systems where operational loads may be exceeded by jamming, ground gusts, control inertia, or friction.

154. § 27.397 (Amendment 27-11) LIMIT PILOT FORCES AND TORQUES.

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

(1) Primary controls, pilot and copilot, should be designed for the limit pilot forces specified in Paragraph (a) of the rule unless higher torces are used.

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

b. Procedures.

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

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

155. § 27.399 DUAL CONTROL SYSTEM.

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

b. Procedures.

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

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

156. § 27.401 (Amendment 27-3) AUXILIARY ROTOR ASSEMBLIES.

a. Explanation.

(1) For rotorcraft equipped with auxiliary rotors, normally called tail rotors, an endurance test is required by § 27.923, and structural strength substantiation is required. Section 27.401(b) specifically refers to structural strength substantiation of detachable blade systems for centrifugal loads resulting from maximum design rotor RPM.

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

Sufficient stiffness and rigidity must be designed into the blades to prevent excessive deformation and to ensure that the blades will maintain the desired aerodynamic characteristics. As a design objective, the structural strength requirements should be met with the minimum material. Excess blade weight imposes extra centrifugal loads that may increase the operating stress levels. Blade weight and strength should be Par 154 205

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optimized. Even though a structural strength analysis for the blade design loads is required, a flight load survey and fatigue analysis are also required by § 27.571.

(3) Section 27.1509 defines the design rotor speed as that providing a 5 percent margin beyond the rotor operating speed limits.

b. Procedures.

(1) The endurance tests prescribed by §§ 27.923 and 27.927 require achieving certain speeds, power, and control displacement for the auxiliary (tail) rotor as well as the main rotor. The parts must be serviceable at the conclusion of the tests.

(2) Structural substantiation of the auxiliary (tail) rotor is required to ensure integrity for the minimum and maximum design rotor speeds and the maximum design rotor thrust in the positive and negative direction. Thrust capability of the rotor should offset the main rotor torque at maximum power as required by § 27.927 (b).

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

The rotor operating speed limits shown during the official FAA/AUTHORITY flight tests must include the noted 5 percent margin with respect to the design speeds.

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

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

(iv) Bearings are generally used in the tail rotor installation to allow flapping and feathering motion of the blades. The bearing manufacturer's ratings of these bearings must not be exceeded. Bearings generally used in main and tail rotors are classified as ABEC Class 3, 5, or 7. Class 7 is the highest quality presently available. Satisfactory completion of the endurance tests of §§ 27.923 and 27.927 is a means of proving that use of a particular bearing is satisfactory.

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(v) The analysis must include appropriate special factors, casting factors, bearing factors, and fitting factors prescribed by §§ 27.619, 27.621, 27.623, and

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

156A. § 27.401 (Amendment 27-27) AUXILIARY ROTOR ASSEMBLIES.

a. Explanation. Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, and 27.341.

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

157. § 27.403 AUXILIARY ROTOR ATTACHMENT STRUCTURE.

a. Explanation.

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

(1) Inertia loads generated by linear and angular accelerations of the auxiliary rotors and their gearboxes, combined with--

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

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

(2) General rules related to proof of structure loads and factor of safety are §§ 27.307, 27.301, 27.303, and 27.305.

b. Procedures.

(1) The angular and linear 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.

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

Additional information can be found in the U.S. Army Material Command Report AMCP 706-201, "Engineering Design Handbook: Helicopter Engineering, Part One, Preliminary Designi 157A. § 27.403 (Amendment 27-27) AUXILIARY ROTOR ATTACHMENT STRUCTURE.

a. Explanation. Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, 27.341.

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

158. § 27.411 GROUND CLEARANCE: TAIL ROTOR GUARD.

a. Explanation.

(1) The rule requires specific protection to prevent the tail rotor from contacting the landing surface during a normal landing if it is possible that the tail rotor will contact the surface. The rule states that it must be impossible for the tail rotor to contact the surface during a normal landing.

(2) If a guard is required, the guard and its supporting structure must withstand suitable design loads.

(3) Section 27.501(c)(1) contains skid landing gear drag requirements that may be applied to the guard design loads.

b. Procedures.

(1) The applicant may submit sketches or drawings showing probable clearance with typical level landing surfaces during normal landings. Typical attitudes such as nose-high autorotation, or autorotation with power-on landing, or other possible tail-low attitudes should be investigated. If the drawings or sketches reveal that it is not likely the tail rotor will contact the landing surface, this minimum clearance with the landing surface may be confirmed during official FAA/AUTHORITY flight tests, such as HV and landing tests. The clearance may be confirmed by having a frangible device of suitable length (i.e., a balsa wood dowel) extending beyond the guard and attached to the tail rotor guard or other appropriate fuselage part. If the device is not damaged, broken, or no contact is made with the surface, compliance has been demonstrated.

(2) If it is possible for the tail rotor guard to contact the landing surface, suitable design loads must be established for the guard. ANC-2a dated March 1948, "ANC Bulletin Ground Loads," Paragraph 6.4, entitled "Tail Bumper Criteria," is an acceptable 208 Par 157

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means of deriving the rotorcraft kinetic energy that shall be absorbed by the guard.

This method is noted here for convenience.

The tail rotor guard shall be able to absorb the kinetic energy of the rotorcraft in its most unfavorable CG position in the tail-down landing attitude. The kinetic energy that the tail rotor guard should be capable of absorbing may be determined by the following:

WV≤ KE = 2g (Ky +133) where-- Vs = vertical speed fusec, derived from § 27.725(a) Ky = pitching radius of gyration - ft from pitching axis 1B = distance from most critical CG location to the guard or bumper contact point - ft g = 32.2 ft/sec W= gross weight less rotor lift from § 27.473(a) - Ibs

(il) Other, more recent, analytical techniques (most utilizing computer programs) may, of course, be used rather than the ANC-2a means after proper substantiation for applicability and validity.

(iii) The tail rotor guard should not fail when the limit and ultimate load, which is derived from a combination of the limit kinetic energy and the guard resulting limit deflection required to dissipate the energy, is imposed on the guard and the rotorcraft tail (see § 27.305).

(3) Substantiation of the guard, skid, or bumper for the design loads derived may be accomplished by test or analysis as stated in § 27.307(a).

(4) Several rotorcraft tail rotor guards are installed solely for the protection of ground personnel from the rotating tail rotor. For guards installed for this purpose, the applicant should use prudent and reasonable design loads and features. Such guards should not present a hazard to the rotorcraft because of its design features.

159. § 27.413 STABILIZING AND CONTROL SURFACES.

a. Explanation. Minimum design loads are specified for stabilizing as well as control surfaces.

(1) Paragraph (a) of the rule requires application of minimum empirical design loads, application of critical maneuvering loads, and application of critical maneuvering loads combined with vertical gust loads (30 feet per second or 17.8 knots per § 27.341).

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(2) Paragraph (b) requires load distributions that closely simulate actual pressure distributions. Both spanwise and chordwise distributions are intended.

(3) These surfaces are used for stability and control thereby hopefully extending the CG range and increasing the airspeed of modern designs.

(4) To "closely simulate actual pressure condition" on the surfaces, unsymmetrical loads are also required on horizontal surfaces. An arbitrary distribution, if conservative, may be used.

(5) It is noted § 27.571 requires fatigue substantiation of the flight structure which will include control and stabilizing surfaces.

(6) If the surface is controllable, a proof and operation test of the surface control system is required by §§ 27.681 and 27.683.

b. Procedures. Modern rotorcraft designs have generally employed a fixed or a wholly movable, not split or divided, stabilizing or control surface.

(1) Design Loads.

Limit loads of 15 pounds per square foot will apply up to approximately a 90-knot design airspeed. Above a 90-knot design airspeed (1.11VNE), the coefficient (Cy = 0.55) imposes higher limit loads on the surface. The coefficient ON is assumed normal to the chordline of the section.

(Ii) In addition, combined maneuvering and vertical up or down gust loads may impose the highest limit loads on the control surfaces of rotorcraft. This is attributed to the change in angle of attack and change in resultant airspeed.

(ili) The applicant may choose to derive the limit loads using maximum aerodynamic coefficients for the surface under consideration at the maximum design airspeed combined with a 17.8-knot gust. This would be acceptable provided these design loads exceed the minimum loads derived from a ON = 0.55 at design airspeed or exceed 15 pounds per square foot load on the surface.

(2) The load distribution on the surface should closely simulate actual pressure distributions.

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(i) The spanwise load may be rectangular, or other acceptable conservative distributions may be used. The method developed by O. Schrenk in NACA TM 948, 1940, is an acceptable method for approximation of spanwise distribution.

NOTE: The method is valid for aspect ratios of 5 through 12 and for rectangular planforms such as used on rotorcraft, other planforms may be acceptable as prescribed in the TM.

(ii) The chordwise distribution appropriate for the aerodynamic shape or a conservative distribution should be used.

(iii) The flight load survey conducted under § 27.571 may be used to confirm design parameters and possible load distribution data. On controllable surfaces, the pitching moment (control loads) may be measured for fatigue substantiation of the control system. The control stabilizing surfaces may be subject to loads measurement and possible fatigue tests for fatigue substantiation also.

(3) Proof of the structure for the required loads is specified in §§ 27.301,

27.303, 27.305, and 27.307. Tests or analysis may be used as prescribed. If analysis is used, fitting factors and other appropriate factors prescribed by the rules of §§ 27.625, 27.621, and 27.623 will be required in the analysis.

159A. § 27.413 (Amendment 27-27) STABILIZING AND CONTROL SURFACES.

a. Explanation. Amendment 27-27 removed this section since the requirements are adequately addressed in §§ 27.337, 27.339, and 27.341.

b. Procedures. The policy material pertaining to this section is retained as supplemental information especially as reference material for Paragraph 139 (§ 27.341) of this document.

160. § 27.427 (Amendment 27-27) UNSYMMETRICAL LOADS.

a. Explanation. Amendment 27-26 added the standard and Amendment 27-27 amended it. Minimum unsymmetrical design loads are specified for horizontal tail surfaces and also vertical tail surfaces whenever they support the horizontal tail surfaces.

(1) Loads are derived by rational analysis, or for earlier certification bases, the prescribed empirical loads of § 27.413 may be used. Section 27.413 was removed by Amendment 27-27 since the requirements are adequately addressed in §§ 27.337,

27.339, and 27.341.

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(2) Rational loads, appropriate for the aerodynamic surfaces, should be distributed according to the standard.

(3) When vertical tail surfaces support the horizontal tail surfaces, the vertical tail surfaces and supporting surfaces are required to support the critical combination of vertical and horizontal surface loads.

b. Procedures. Two basic loading conditions are required by § 27.427 for each of the two basic empennage configurations.

(1) Horizontal surfaces supported by the tail boom or fuselage. Structural substantiation should be provided for all six combinations shown in Figure 160-1. All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly.

First unsymmetrical loading condition:

(A) 100 percent of the flight load is applied to one side of the plane of symmetry; and 0 percent of the flight load is applied to the other side of the plane of symmetry.

(B) For surfaces with end plates or other similar devices, the load distribution will be changed accordingly.

(i) Second unsymmetrical loading condition:

50 percent of the flight load is applied to one side of the plane of symmetry acting up;

and 50 percent of the flight load is applied to the other side of the plane of symmetry acting down.

(2) Horizontal surfaces supported by a vertical surface. Structural substantiation should be provided for all six combinations shown in Figure 160-2. All of these empirical loading distributions should be used unless rational analysis shows one or more of each set of conditions to be non-critical or equal or more realistic distributions are substantiated. Rectangular spanwise air load distribution should be used unless more rational distribution is substantiated. If end plates are used, the air loads should be distributed accordingly.

First unsymmetrical loading condition:

100 percent of the flight load is applied to one side of the plane of symmetry; and O percent of the flight load is applied to the other side of the plane of symmetry.

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(ii) Second unsymmetrical loading condition:

50 percent of the flight load is applied to one side of the plane of symmetry acting up;

and 50 percent of the flight load is applied to the other side of the plane of symmetry acting down.

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Figure 160-1. (View Looking Forward) 214 Par 160

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Figure 160-2. (View Looking Forward) Par 160 215

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**161.-169. RESERVED.**

Intentionally Left Blank 216 (thru 224) Par 161

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**SECTION 10. GROUND LOADS**

170. $ 27.471 GENERAL.

a. Explanation. This regulation specifies that limit ground loads must be considered which are:

(1) External loads caused by landing (ground) conditions for skid and wheel landing gear equipped rotorcraft and by ground taxiing loads as specified in § 27.235 for wheel landing gear equipped rotorcraft.

(2) Loads considering the rotorcraft structure as a rigid body.

(3) Loads in equilibrium with linear and angular inertia loads.

(4) The critical center of gravity "must be selected so that the maximum design loads are obtained in each landing gear element."

b. Procedures.

(1) The standards to be considered are specified in §§ 27.473 through 27.505.

These associated standards cover landing gear arrangements, landing conditions, and ground loading conditions (for wheel landing gear rotorcraft).

(2) Drop tests may be used to verify landing load factors. (See Paragraph 299 of this document.)

(3) The application of the design loads derived from the landing load factors will be as specified for each element affected by landing or ground loading conditions (for wheel landing gear rotorcraft).

(4) During the applicant's flight test program, the landing load factors for skid and wheel landing gear rotorcraft and taxiing load factors for wheel landing gear rotorcraft are monitored to assure the design load factors used are adequate. See Paragraph 97 of this document for § 27.235 policy.

171. § 27.473 (Amendment 27-2) GROUND LOADING CONDITIONS AND ASSUMPTIONS.

a. Explanation. The rotorcraft is to be designed for the maximum weight. A rotor lift of two-thirds of the design maximum weight may be used. The minimum limit landing load factor is determined by the drop tests of § 27.725.

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b. Procedures. Loads for the landing conditions are derived considering mass (equal to the maximum weight) and rotor lift (equal to two-thirds of the maximum weight) acting through the center of gravity throughout the landing impact. Unbalanced external loads resulting from asymmetric loading conditions are reacted as specified in the individual subparagraphs. The rotorcraft must be substantiated for ultimate landing loads by either test or analysis utilizing an ultimate load factor of 1.5 applied to the limit load factor of not less than that substantiated under § 27.725.

172. § 27.475 TIRES AND SHOCK ABSORBERS.

a. Explanation. This section specifies the tire and shock absorber position to be used in ground load derivations.

b. Procedures. Ground loads are to be derived with the tires in static (1g) position and the shock absorbers "in their most critical position." The determination of the "most critical positioni for the shock absorbers generally requires a load versus deflection test or analysis of the shock absorber system and a determination of the effect of both load and deflections on the shock absorber, attachment structure, and substructure designed by ground loads.

173. § 27.477 LANDING GEAR ARRANGEMENT.

a. Explanation. This section specifies the individual standards to be used for ground load conditions for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity.

NOTE: § 27.497 gives ground loading conditions for landing gear with tail wheels, and § 27.501 gives ground loading conditions for landing gear with skids.

b. Procedures. The ground loading conditions of §§ 27.235, 27.479 through

27.485, and 27.493 will be used for rotorcraft having two wheels aft and one or more wheels forward of the center of gravity. This includes forward wheels on separate axles.

174. § 27.479 LEVEL LANDING CONDITIONS.

a. Explanation. This section provides explicit level landing load criteria for landing gear with two wheels aft and one or more wheels forward of the center of gravity.

(1) Level landings--

(i) Each wheel contacting the ground simultaneously; and

(ii) Aft wheels contacting the ground with forward wheels just clear of the ground.

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(2) Application of loads-- Maximum design vertical loads applied alone; and

(i) The maximum design vertical loads applied with a drag load of at least 25 percent of the vertical load (applied at the ground contact area).

(3) A 40 percent/60 percent load distribution between wheels for configurations having two forward wheels including quadricycle. This distribution between wheels on a common axis is to be applied for the conditions of vertical loads only and for vertical loads combined with drag loads of 25 percent of the vertical loads.

(4) Aircraft pitching moments are to be reacted by the forward landing gear for simultaneous wheel contact or by the angular inertia forces when the forward landing gear is clear of the ground as specified.

b. Procedures.

(1) The specified loading conditions will be used in load derivations.

(2) The critical center of gravity condition will be used for each gear and gear support structure.

(i) The aft center of gravity condition with the forward gear clear will normally be critical for the aft gear and gear supports.

(ii) The forward center of gravity condition with each gear contacting the ground simultaneously will normally design forward gear elements critical for vertical loads.

(iii) The forward center of gravity condition with the forward gear clear may result in high load factors, angular plus linear, that will greatly affect security of items of significant mass.

175. § 27.481 TAIL-DOWN LANDING CONDITIONS.

a. Explanation. This section provides the criteria for tail-down landing conditions;

i.e., "the maximum nose-up attitude allowing ground clearance" with ground loads acting "perpendicular to the ground."

b. Procedures.

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(1) The tail-down landing condition will be used to check (by analysis or test) for criticality of landing gear or support structure. This attitude generally creates the highest forward loads on the main landing gear in combination with vertical loads.

(2) The tail-down landing condition may be the critical condition for both landing load factor and for energy absorption by the main gear. Section 27.725 requires that "each landing gear must be tested in the attitude simulating the landing condition that is most critical." Where questions exist as to the critical attitude, both level landing and tail-down landing attitudes should be used in drop tests required by § 27.725.

176. § 27.483 ONE-WHEEL LANDING CONDITIONS.

a. Explanation. This section gives the condition to be used for one-wheel landing conditions. Only the vertical load condition of § 27.479(b)(1) is required.

b. Procedures. The one-wheel landing condition is generally critical for the landing gear-to-fuselage attachments and the landing gear elements between the attachments. Unbalanced external loads are reacted by rotorcraft inertia.

177. § 27.485 LATERAL DRIFT LANDING CONDITIONS.

a. Explanation.

(1) This section provides the loading conditions which impose side (and vertical) loads on the landing gear. A level landing attitude is specified. Two main conditions required are--

(i) Only the aft wheels in contact with the ground; and

(ii) All wheels contacting the ground simultaneously.

(2) Loads. The vertical loads to be applied with the side loads are specified as "one-half of the maximum ground reactions of § 27.479(b)(1)." These vertical loads are the level landing loads considering both contact and noncontact with the ground by the forward wheels.

(i) One side load condition is specified as "0.8 times the vertical reaction acting inward on one side and 0.6 times the vertical reaction acting outward on the other side" when only the aft wheels contact the ground.

(ii) The other side load condition (for all wheels contacting the ground) specifies the 80 percent inward/60 percent outward distribution for the aft wheels and

0.8 times (80 percent) the vertical reaction for the forward wheels.

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b. Procedures. The loading conditions, as specified, are applied to the landing gear and attaching structure. The loads are applied at the ground contact point, except for full swiveling gear which has the load applied at the center of the axle. In other words, full swiveling gear is considered to have swiveled to a static position under the side load before the design vertical and side loads are achieved. The rotorcraft as well as the landing gear itself will be substantiated for these side load conditions.

178. § 27.493 BRAKED ROLL CONDITIONS.

a. Explanation. This section provides two loading conditions for ground braking operations. Specific vertical loads in conjunction with drag loads (due to braking) are to be considered. The limit vertical load factor is 1.33 for condition of all wheels in contact with the ground and 1.0 for condition of aft wheels only in contact with the ground and nose wheel clear. The drag load on wheels with brakes is 0.8 times the vertical load or the drag load value based on limiting brake torque, whichever is less. The drag load value for limiting brake torque may be that determined in the performance testing to TSO C26 or equivalent, as required.

b. Procedures. The braking loads are calculated from the specified criteria with the shock absorbers in their static (normal) positions and with the drag loads applied at the ground contact point. Structural substantiation of the affected structure may be accomplished by test or analysis. If tests are used, the wheel and tire assembly is commonly replaced with a test fixture so the limit loads and static deflections specified can be more accurately controlled. The test specimen should be complete enough to ensure that the landing gear structure and the attach and backup structure are adequately substantiated.

179. § 27.497 GROUND LOADING CONDITIONS: LANDING GEAR WITH TAIL WHEELS.

a. Explanation. This section provides the loading conditions for landing gear designs with tail wheels.

(1) Level landings are to consider the following:

(i> All wheels (main and tail) contacting the ground simultaneously, as well as only forward main wheels contacting the ground.

(ii) Maximum design vertical loads applied alone.

(iii) The maximum design vertical loads combined with a drag load of at least 25 percent of the vertical loads for both conditions.

(2) Noseup landings with only the rear wheel or wheels initially contacting the ground must be considered unless shown to be extremely remote.

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(3) Level landings on one forward wheel only are to be considered. Drag loads are not required.

(4) Side load conditions are imposed on the main wheels and tail wheels for level landing attitudes. Criteria for full swiveling and locked tail wheels are included in this standard.

(5) Braked roll conditions are specified for the level landing attitudes.

(6) Rear wheel turning loads are also specified for swiveling and locked tail wheels.

(7) Taxiway condition loads for the landing gear and rotorcraft are those that ioccur when the rotorcraft is taxied over the roughest ground that may reasonably be expected in normal operation." The aircraft design load factors should not be exceeded during the evaluation. Section 27.235 contains an identical standard that applies to all types of wheel landing gear.

b. Procedures.

(1) The specified loading conditions are to be used in load derivations.

(2) The critical center of gravity condition is used for each gear and gear support structure.

The forward center of gravity condition with the tail gear clear will normally be critical for the forward gear and gear supports.

(її) The aft center of gravity condition with the tail gear clear should be checked for criticality of security of large mass items located forward of the center of gravity. Vertical and angular accelerations are additive under this landing condition.

(ili) The aft center of gravity condition with each gear contacting the ground simultaneously will generally design tail gear elements critical for vertical loads.

The other conditions are generally less severe but must be proven.

(3) For tail-down landing procedures use § 27.481. The reference to "extremely remote" in § 29.497(d)(2) predates current §§ 25.1309, 29.1309, and AC 25.1309.1. This phrase has been used to require consideration of noseup landings unless features of design are present which prevent noseup landings or where such landings are unlikely during the life of the rotorcraft. (See Paragraph 175 of this document.)

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(4) Use § 27.483 for one-wheel landing procedures, Paragraph 176 of this document.

(5) Use § 27.485 procedures for side load conditions, Paragraph 177 of this document.

(6) Use § 27.493 procedures for braked roll conditions, Paragraph 178 of this document.

(7) For rear wheel turning loads, swiveling of tail landing gears is allowed as in basic side load conditions. The side load is applied at the axle or, if the wheel is locked, the load is applied at ground contact. Rear wheels are loaded with the critical vertical static load in conjunction with an equal side load to substantiate the tail gear.

(8) Since the rotorcraft is to be designed for load factors that will not be exceeded during taxi tests or other conditions, an instrumented taxi test program will be necessary. (Use § 27.235, Paragraph 97, of this document.)

180. § 27.501 (Amendment 27-2) GROUND LOADING CONDITIONS:

**LANDING GEAR WITH SKIDS.**

a. Explanation. This section provides the ground loading conditions for landing gear with skids. The loading conditions are similar to those for wheeled gear except for the following criteria which are unique to skid gears:

(1) Structural yielding of elastic spring members under limit loads is allowed.

(2) Design ultimate loads for elastic spring members need not exceed the loads obtained in a drop test with a drop height of 1.5 times the limit drop height. The rotorcraft and the landing gear attachments are subject to the prescribed design ultimate loads.

(3) The gear must be in its most critically deflected position (similar to § 27.475).

(4) Ground reactions are rationally distributed along the bottom of the skid unless otherwise specified. Section 27.501(f) concerns specific "concentrated" and arbitrary load conditions.

(5) Drag loads are 50 percent of vertical reactions rather than the 25 percent for wheeled gear.

(6) Side loads are 25 percent of the total vertical reaction rather than the 60 to 80 percent for wheeled gear.

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(7) Side loads are applied to one skid only (inward acting and outward acting) with resulting unbalanced moment resisted by angular acceleration.

(8) A ground reaction load of 1.33 times the maximum weight is to be applied at 45° from the horizontal axis:

Distributed among or between the skids;

(ii) Concentrated at the forward end of the straight portion of the skid tube; and

(ill) Applied only to the forward end of the skid tube and its attachment to the rotorcraft.

(9) A concentrated vertical load equal to one-half of the design limit vertical load is to be applied at a point midway between the skid tube attachments. This condition applies only to the skid tube and its attachment to the rotorcraft.

b. Procedures.

(1) The specified loading conditions are to be used in load derivations.

(2) The critical center of gravity conditions are to be used for each gear and gear support structure. Asymmetry of the skid tubes, cross tubes, and gear attachments is to be considered in determining the critical center of gravity condition.

(3) The rotorcraft and landing gear attachment must be substantiated for ultimate landing loads by either test or analysis utilizing an ultimate load factor of 1.5 in accordance with § 27.303. The elastic spring members may be analyzed or static tested for ultimate loads (and deflections) using either a factor of safety of 1.5 or one associated with an "ultimate" drop height of 1.5 times the limit drop height.

Substantiation by "ultimate" drop tests may be used provided all combinations of critical parameters are included in the total substantiation effort. This method will require a series of tests using several test specimens or a limited number of drop tests plus further substantiations by static tests or analyses for additional critical conditions not covered by the drop tests).

180A. § 27.501 (Amendment 27-26) GROUND LOADING CONDITIONS:

**LANDING GEAR WITH SKIDS.**

a. Explanation. Amendment 27-26 relaxes the previous requirements in two cases by:

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(1) Allowing the total sideload of Paragraph 27.501(d)(3) to be distributed "equally between skids" rather than being "applied along the length of one skid only;"

and,

(2) Allowing the concentrated load of Paragraph 27.501(f)(2) (ii) to be distributed over the central 33.3 percent of the skid (between skid tube attachments) rather than being "concentrated at a point midway between the skid tube attachments."

b. Procedures. The previous procedures (through Amendment 27-19) continue to apply to Amendment 27-26 except use the new load distributions.

181. § 27.505 SKI LANDING CONDITIONS.

a. Explanation. This is an optional requirement for ski operations. The regulation specifies vertical loads, side loads, and torque loads (Mz) to be applied to ski installations. The four loading conditions to be applied at the pedestal bearings are:

(1) Simultaneous application of Pn, up load, and Pn/4, horizontal load.

(2) Up load of 1.33 P.

(3) Side load of 0.35 Pn.

(4) Torque load of 1.33 P (in foot-pounds) about the vertical axis through the centerline of the pedestal bearings.

NOTE: Where P is the maximum static weight on each ski and n is the limit load factor obtained from drop tests. The load factor obtained from wheel or skid landing gear drop tests may be used.

b. Procedures. Structural substantiation may be accomplished by static test or analysis using the specified loads. Skis generally have a limit load rating. The design loads derived for this standard must not exceed the rating. (TSO-C28 concerns, in part, standards for aircraft skis.)

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**182.-192. RESERVED.**

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**SECTION 11. WATER LOADS**

**193. RESERVED.**

194. § 27.521 FLOAT LANDING CONDITIONS.

a. Explanation. This is an optional requirement for float operations. The regulation specifies vertical loads, aft loads, and side loads to be applied to the float installations. The two loading conditions to be applied are:

(1) Up-load Condition.

A vertical load appropriate to a landing load factor determined under § 27.473(b).

(її) A resultant water reaction passes vertically through the aircraft CG.

(III) An aft load equal to 25 percent of the vertical load.

(2) Side-load Condition.

A vertical load equal to 75 percent of the vertical load for the up-load condition.

(її) A vertical load equally divided among the floats.

(ili) A side load at each float equal to 25 percent of the vertical load of Paragraph (a)(2)(i) at each float.

b. Procedures.

(1) The vertical load factor is determined by drop tests in accordance with §§ 27.473(b) and 27.725. The floats may be drop tested, or they may be assumed to have the same load factor as wheeled gear which have been drop tested.

(2) Structural substantiation may be accomplished by either static tests or analysis using the specified loads. The load distribution on the floats may be realistic, based on hydrostatic pressure distributions, or conservative.

**195.-204. RESERVED.**

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**SECTION 12. MAIN COMPONENT REQUIREMENTS**

205. § 27.547 (Amendment 27-3) MAIN ROTOR STRUCTURE.

a. Explanation. This regulation requires the main rotor structure to be designed to the static load requirements of §§ 27.337 through 27.341 (vertical maneuvering loads and vertical and horizontal gust loads). In addition, the main rotor blades, hubs, and flapping hinges are specified to be designed for impact forces of each blade against its stop during ground operation and for specified limit torque at any rotational speed including zero. The torque forces (from the drive system) are distributed to the rotor blades as specified.

b. Procedures.

(1) Substantiation in compliance with this standard is accomplished by application of the flight loads of §§ 27.337 through 27.341 and the torque loads of § 27.361 to the rotor structure by stress analyses and/or static tests. The use of wind tunnel data as well as flight loads survey data may be used to generate and/or check the external load magnitudes and distributions.

(2) Where new materials are used in the main rotor structure, such as composites containing plastics, the effects of temperature and humidity are to be considered in accordance with § 27.603, and the effects of uncertainties in manufacturing processes or inspection methods are to be considered in accordance with § 27.619. More experience is available for metallic materials, but § 27.603 requires that metallics be suitably protected against the effects of environmental conditions.

(3) The design impact forces of each blade must be imposed against its stop or stops. Appropriate monitoring of the blades, hubs, flapping hinges, and stops during laboratory tests, ground endurance tests, and flight tests should ensure that the stops are sufficient for ground operation loads. The design torque loads are derived as prescribed.

206. § 27,549 (Amendment 27-3) FUSELAGE, LANDING GEAR, AND ROTOR PYLON STRUCTURE.

a. Explanation. This regulation requires that the fuselage, landing gear, and rotor pylon (including the tail fin, if any) be designed to withstand the flight loads of §§ 27.337 through 27.351, the ground loads of §§ 27.235, 27.471 through 27.497, skid loads of § 27.501, ski loads of § 27.505, water loads of § 27.521, and rotor loads of § 27.547 (d) and (e). The ski and water loads pertain to optional features. Consideration is also required of -- 258 Par 205

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(1) Auxiliary rotor thrust;

(2) The torque reaction of each rotor drive system; and

(3) Balancing air and inertia loads.

b. Procedures. Compliance with this standard is accomplished by application of the specified aircraft loads including engine torque to the fuselage and rotor pylon structure by stress analyses and/or static tests. Drive system torque factors to be used are noted in § 27.547 for the main rotor structure as well as in Paragraph (e) of this standard.

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**207.-217. RESERVED.**

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**SECTION 13. EMERGENCY LANDING CONDITIONS**

218. § 27.561 GENERAL.

a. Explanation.

(1) The occupants should be protected as prescribed from serious injury during an emergency/minor crash landing on water or land for the conditions prescribed in the standard. The standard states that each occupant should be given every reasonable chance of escaping serious injury in a minor crash landing.

(2) The standard in § 27.561(b)(3) specifies certain ultimate inertia load factors but allows a lesser downward vertical load factor by virtue of a 5 FPS ultimate rate of descent.

(3) In addition, the occupants shall be protected from items of mass inside the cabin as well as outside the cabin. For example, a cabin fire extinguisher must be restrained for the load factors prescribed in this section (reference § 27.1411(b)(2).) A transmission or engine must be restrained to the load factors in § 27.561(b)(3) if located above or behind the occupants.

(4) For aircraft equipped with retractable landing gear, the landing gear is retracted for compliance with the standards.

b. Procedures.

(1) The design criteria report or another similar report of the rotorcraft structural limits should contain the (ultimate) minor crash condition load factors.

(2) Section 27.785 (Paragraph 336 of this AC) concerns application of this design standard to seats, berths, belts, and harnesses.

(3) The ultimate design landing and maneuvering load factors may exceed the minor crash condition load factors. The highest load factor derived shall be used.

For example, for light weight conditions, the ultimate maneuvering load factor may be 5.25g as specified in § 27.337.

(il) The ultimate vertical landing load tactors derived trom 99 27.471 through 27.521, whichever is appropriate for the design, may exceed the 4.0g down load factor in this section. The rotorcraft landing case design limit contact velocity shall range from 6.5 to 8.3 FPS (see §§ 27.473 and 27.725).

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(4) As specified in (b)(3) (iv) of the standard, the downward load factor is 4.0 or a lower design load factor may be used.

The lower load factor relates to a rotorcraft impacting a flat, hard landing surface at 5 FPS (ultimate) vertical rate of descent. The load factor derived for each unique design is a tunction of the aircraft impact/crushing characteristics.

(i) The 4.0g down load factor case is related to either a fixed or retractable gear rotorcraft. This condition is not dependent on impact characteristics of the rotorcraft.

(ill) As noted in Paragraph (3) above, the design landing load factors may exceed each of the two previous cases and would then become the prominent design (vertical load) parameter for seats, transmissions, fire extinguishers, etc.

(5) Items of mass such as fire extinguishers, nav-com equipment, liferafts, engines, and/or transmissions shall be restrained for the appropriate load factors.

(6) Cargo/baggage compartments separated from the passenger compartment shall be designed for load factors specified in § 27.787. The conditions in § 27.561 are excepted from that standard.

218A. § 27.561 (Amendment 27-25) EMERGENCY LANDING CONDITIONS - GENERAL.

a. Explanation. Amendment 27-25 adds or increases the design static load factor of § 27.561 in two areas:

(1) The design static load factors for the cabin in § 27.561(b)(3) are increased in concert with the dynamic test requirements of new § 27.562.

(2) Design static load factors are added in § 27.561(c) for external items of mass located above and/or behind the crew and passenger compartment.

b. Procedures. The procedures of Paragraph 218, § 27.561, continue to apply except the new load factors of § 27.561 should be used. Penetration of any items of mass into the cabin or occupied areas should be prevented.

218B. § 27.561 (Amendment 27-30) EMERGENCY LANDING

**CONDITIONS - GENERAL**

a. Explanation.

(1) Amendment 27-30 adds Paragraph (d) which lists specific load factors for the fuselage structure in the area of internal fuel tanks located below the passenger 272 Par 218

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floor level. For other locations, the fuselage structure is to be designed to resist crash impact loads prescribed in § 27.561(b)(3) for fuel tanks located within the cabin area; or § 27.561(c) for fuel tanks located behind or above the occupant area. These load factors are provided to prevent crash induced fuel tank ballistic hazards to occupants and to also protect the fuel tank from rupture as prescribed. The landing gear must be retracted if the rotorcraft is equipped with retractable gear.

(2) Section 27.952(b) provides specific load factors for the fuel tanks which are identical to the load factors stated in § 27.561. Paragraph 447 of this document provides information and guidance for § 27.952 and may be used in conjunction with this paragraph.

b. Procedures. The procedures of Paragraphs 218 and 218A continue to apply except new load factors are established for fuel tanks located below the passenger floor level. Each fuel tank and its installation are subject to the loads stated in the standard.

The load factors are determined by the fuel tank location.

218C. § 27,561 (Amendment 27-32) EMERGENCY LANDING CONDITIONS - GENERAL.

a. Explanation. Amendment 27-32 adds a new rearward emergency load factor of 1.5g to both §§ 27.561(b)(3)(v) and 27.561(c)(5). The addition of the 1.5g rearward load factor in § 27.561(b)(3)(v) is to provide an aft ultimate load condition for substantiation of the restraints required for retention of both occupants and significant items of mass inside the cabin that could otherwise come loose and cause injuries in an emergency landing. The addition of the 1.5g rearward load factor to § 27.561(c)(5) is to provide an aft ultimate load condition for substantiation of the support structure for retention of significant items of mass above and forward of the occupied volume(s) of the rotorcraft that could otherwise come loose and injure an occupant in an emergency landing. Amendment 27-32 also increases the forward, sideward, and downward emergency load factors of § 27.561(c)(2), (c)(3), and (c)(4), respectively; for retention of items of mass above and behind the occupied volumes) that could otherwise come loose and injure an occupant in an emergency landing.

b. Procedures. The procedures of Paragraphs 218, 218A, and 218B continue to apply except the newly specified load factors must be used. A list of the significant items of mass to be considered should be compiled by the applicant and approved by the certifying authority.

219. § 27.562 (Amendment 27-25) EMERGENCY LANDING DYNAMIC CONDITIONS.

a. Explanation. Amendment 27-25 adds new requirements for the dynamic testing of all seats in rotorcraft.

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b. Procedures. AC 20-137, "Dynamic Evaluation of Seat Restraint Systems and Occupant Restraint for Rotorcraft (Normal and Transport)i provides procedures for complying with § 27. 562 using the 170-pound anthropomorphic test dummy specified in § 27.562(b). Those seats not occupied for takeoff and landing, and so placarded and identified in the rotorcraft flight manual (RFM), may be excluded from compliance.

220. § 27.563 (Amendment 27-11) STRUCTURAL DITCHING PROVISIONS.

a. Explanation. Amendment 27-11 included certification requirements for ditching approvals. The rotorcraft must be able to sustain an emergency landing in water as prescribed by § 27.801(e).

b. Procedures. Refer to Paragraph 338 of this AC for procedures.

220A. § 27.563 (Amendment 27-26) STRUCTURAL DITCHING PROVISIONS.

a. Explanation. Amendment 27-26 added specific structural conditions to be considered to support the overall ditching requirements of § 27.801. These conditions are to be applied to rotorcraft for which over-water operations and associated ditching approvals are requested.

(1) The forward speed landing conditions are specified as:

(i) The rotorcraft should contact the most critical wave for probable water conditions, in the likely pitch, roll, and yaw attitudes.

(її) The forward velocity relative to wave surface should be in a range of 0 to 30 knots with a vertical descent rate of not less than 5 FPS relative to the mean water surface.

NOTE: A forward velocity of less than 30 knots may be used for multiengine rotorcraft if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out touchdown.

(iii) Rotor lift of not more than two-thirds of the design maximum weight may be used to act through the CG throughout the landing impact.

(2) For floats fixed or deployed before water contact, the auxiliary or emergency float conditions are specified in § 27.563(b)(1). Loads for a fully immersed float should be applied (unless it is shown that full immersion is unlikely). If full immersion is unlikely, loads resulting from restoring moments are specified for sidewind and unsymmetrical rotorcraft landing.

(3) Floats deployed after water contact are normally considered fully immersed during and after full inflation. An exception would be when the inflation interval is long 274 Par 219

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enough that full immersion of the inflated floats does not occur; e.g., deceleration of the rotorcraft during water impact and natural buoyancy of the hull prevent full immersion loads on the fully inflated floats.

b. Procedures.

(1) The rotorcraft support structure, structure-float attachments, and floats should be substantiated for rational limit and ultimate ditching loads.

(2) The most severe wave heights for which approval is desired are to be considered. A minimum of Sea State 4 condition wave heights should be considered (reference Paragraph 338 (§ 27.801) of this AC for a description of Sea State 4 conditions).

(3) The landing structural design consideration should be based on water impact with a rotor lift of not more than two-thirds of the maximum design weight acting through the center of gravity under the following conditions:

(i) Forward velocities of 0 to 30 knots (or a reduced maximum forward velocity if it can be demonstrated that a lower maximum velocity would not be exceeded in a normal one-engine-out landing).

The rotorcraft pitch attitude that would reasonably be expected to occur in service. Autorotation flight tests or one-engine-inoperative flight tests, as applicable, should be used to confirm the attitude selected. This information should be included in the Type Inspection Report.

(III) Likely roll and yaw attitudes.

(iv) Vertical descent velocity of 5 FPS or greater.

(4) Landing load factors and water load distribution may be determined by water drop tests or analysis based on tests.

(5) Auxiliary or emergency float loads should be determined by full immersion or the use of restoring moments required to react upsetting moments caused by sidewind, asymmetrical rotorcraft landing, water wave action, rotorcraft inertia, and probable structure damage and punctures considered under § 27.801. Auxiliary or emergency float loads may be determined by tests or analysis based on tests.

(6) Floats deployed after initial water contact are required to be substantiated by tests or analysis for the specified immersion loads (same as for (5) above and for the specified combined vertical and drag loads).

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**221.-229. RESERVED.**

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**SECTION 14. FATIGUE EVALUATION**

230. § 27.571 (Amendment 27-26) FATIGUE EVALUATION OF FLIGHT STRUCTURE.

a. Explanation. An evaluation is required to assure structural reliability of the rotorcraft in flight.

(1) Advisory Circular 20-95 contains background information and acceptable means of compliance with the requirements. A safe life may be assigned or the structure may be fail safe as prescribed or a combination of these may be used.

(2) Mandatory inspections, service life (replacement times) etc., determined in complying with the standard shall be placed in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness (also called Maintenance Manual). See Appendix A of FAR Part 27, Paragraphs A27.4 and Paragraph 729 of this document for information.

(3) Amendment 27-26 amended the standard to require evaluation of the landing gear and their related primary attachments.

(4) Amendment 27-26 also amended the standard to require evaluation of ground-air-ground cycles on the rotorcraft, and if applicable, of external cargo operations. Previously external cargo operations were evaluated whenever the rotorcraft cargo combination exceeded the "standard" maximum certificated gross weight, and the CG range specified in § 27.25(c). If these limits were not exceeded, an evaluation was not required by the standard prior to Amendment 27-26.

b. Procedures.

(1) The fatigue evaluation requires consideration of the following factors:

Identification of the structure/components to be considered.

(ii) The stress during operating conditions.

(Iii) The operating spectrum or frequency of occurrence including frequency of ground-air-ground cycles, as well as external cargo operations.

(iv) Fatigue strength, and/or fatigue crack propagation characteristics, residual strength of the cracked structure.

(2) Since the design limits, e.g., rotor RPM (maximum and minimum), airspeed, and blade angles (thrust, weight, etc.) affect the fatigue life of the rotor system, it is Par 230 285

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necessary that flight conditions be conducted at limits that are appropriate for the particular rotorcraft and at the correct combination of these limits. It will be the responsibility of engineering and flight test personnel to determine that the flight strain program proposal includes conditions of flight at the various combinations of rotor RPM, airspeed, thrust, etc., that will be representative of the limits used in service. The flight test personnel should assure that the severity of the maneuvers to be investigated is such that actual service use will not be more severe. Verification that proposed maneuvers are suitable may be achieved by:

(i) Flying a representative set of maneuvers with the applicant's pilot in the test aircraft at noncritical combinations of weight, CG, and speed. (An FAA/AUTHORITY letter for specific test authorization would ordinarily be required.) If the procedure is used, the applicant should provide adequate preliminary flight strain data from development or other tests to confirm a cleared (non-critical) flight envelope for conduct of these representative maneuvers.

(ii) Flying a representative set of maneuvers with the applicant's pilot in a similar (certified) model to assess and agree upon the required maneuvers, control deflections, and aircraft rates. The required maneuvers or conditions will be specified in the flight strain program plan.

(iii) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.

(iv) Observation of telemetered flight data to assure desired control deflections, rates, and aircraft attitudes.

(V) Some combinations of items b(2)(i) through b(2) (iv) above.

(3) Assessing the operation spectrum and the flight loads or strain measurement program will involve airframe, propulsion, and flight test personnel.

(4) Variation in the operating or loading spectrum among models, and variations in the spectrum for a particular model rotorcraft, should be evaluated.

AC 20-95, Paragraph 7, entitled "Loading Spectrum," contains the statement that Table 1 (of the circular) contains typical percent of occurrences for various flight conditions for a single-piston-engine powered small rotorcraft used in utility operations.

In addition, the table should be used only as a guide and should be modified as necessary for each particular rotorcraft design.

(5) The difference in loading spectrum for different models that may be anticipated is illustrated by comparing the percentage of time assigned to level flight conditions, specifically 0.8 Vu to 1.0 Vu for three different rotorcraft designs as shown in Table 230-1. (V. is the maximum airspeed at maximum continuous power in level flight.) The first column was obtained from Table 1, AC 20-95 which applies to a 286 Par 230

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single-piston-engine powered small rotorcraft used in utility operations. The second column is appropriate tor a single-turbine-engine powered seven-place small business and utility rotorcraft. The third column is appropriate for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in Table 230-1 for the turbine utility business and turbine transport rotorcraft are examples of particular designs. The high percentage of time shown in this level flight regime could be unconservative for some designs, especially if the stresses under these design conditions produce an infinite fatigue life for the particular component. The fatigue spectrum percentage of occurrences in AC 20-95 shall be modified according to the intended operational usage of the rotorcraft. However, a conservative application should be considered. This variation illustrates the "tailoring" of the loading spectrum for the type of rotorcraft and the anticipated usage.

Table 230-1 Comparison Percent of Time in Level Flight Turbine Piston Utility Twin Turbine Utility Business Transport

0.8 VNE 25% 0.8 VH 16% 0.8 VH 15%

1.0 VH 15% 0.9 VH 21% 0.9 VH 20%

1.0 VNE 3% 1.0 VH 24% 1.0 VH 38% Total 43% 61% 73%

(6) External cargo operations are a unique and demanding operation. A "logging" operator may use 50 maximum power applications per flight hour to move logs from a cutting site to a hauling site. Power is used to accelerate, decelerate, or hover prior to load release. Lifting loads over an obstruction or natural barrier is another example of very frequent high power applications for takeoff and for hovering over the release area. Similar types of operations require flight loads data to assess the effects on fatigue critical components.

(7) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives, inspections, etc., will be affected. The assessment may be done by calculating an "external cargo configuration" service life for each critical component. The lowest service life obtained from standard configuration flight loads data and loading spectrum, or from external cargo configuration flight loads data and loading spectrum or from frequent ground-air-ground cycles is generally the approved service life or replacement time. Since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations, this procedure could be used if an "operational cycles" equation for equivalent flight hours is not approved (see (8) below).

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(8) The Airworthiness Limitations Section of the maintenance manual shall contain the required information derived from complying with the standard. If an "operational cycles" equation for "equivalent flight hours" is approved under the standard, the equation is included in this approved section of the manual.

(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if applicable. The ground-air-ground cycle is inherent in these conditions. This procedure will avoid delays associated with reinstallation and calibration of equipment.

230A. $27.571 (Amendment 27-33) FATIGUE EVALUATION OF FLIGHT STRUCTURE FOR CATEGORY A CERTIFICATION.

a. Explanation. Amendment 27-33 added Appendix C to specify the requirements for Category A certification of normal category rotorcraft. The requirement for fatigue tolerance evaluation will require test evidence to support the analysis.

b. Procedures. For Category A certification, the tests specified in AC 29-2A, Section 230A are required for fatigue tolerance evaluation. AC 29-2A, Section 230A is repeated in this section.

(1) Fatigue test evidence is necessary for the fatigue evaluation of gears. The test evidence should be provided by rotating tests of complete gearbox specimens operating under power. The tests provide the basis for analysis leading to the establishment of safe life.

(2) The tests are conducted specifically for the purpose of gear tooth evaluation, and components subjected to the tests do not have to be considered serviceable on completion of the test. Excessive wear on bearings and shafts and marking (including spalling) of bearings and gear teeth are acceptable provided no fatigue damage is evident on the gear teeth. However fatigue damage other than tooth fatigue should be considered for test validity and the integrity of the affected part confirmed as necessary.

(3) The test conditions (torque versus number of cycles) should permit the setting of mean strength curve(s) to be associated with each primary gear in the drive train. The test conditions, should at a minimum, encompass those power levels for which repeated application inservice is expected under normal circumstances. The S-N curve(s), for the material and type of gear, should be reduced by a factor of safety to take into account material and manufacturing variability. The factored curve will then be used in conjunction with the flight power spectrum to determine a life (limited or unlimited) for the gears in the primary drive system.

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(4) Special procedures, which do not affect fatigue evaluation of the gear teeth, may be allowed to facilitate completion of the test provided they have been justified and they do not affect life determination. These include periodic interruption for inspection, etc., replacement of non-critical parts and the use of special lubricants, special cooling systems, and methods to prevent unrepresentative deflections at the test torque levels.

(5) From evidence in relation to the strength of steel gears of conventional design, it is accepted that adequate fatigue strength can be demonstrated by the use of the above safety factor of 1.4 for a single test, 1.35 for two tests, 1.32 for three tests, and 1.3 for four or more tests. Where several tests are to be conducted, specimens should be selected from different manufacturing batches if practicable.

(6) The demonstration of infinite life for gear teeth will normally require tests of a minimum of 107 cycles duration at factored power levels. Use of shorter duration tests should be justified.

**231.-240. RESERVED.**

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

241. $ 27.601 DESIGN.

a. Explanation.

(1) This rule requires that no design features or details be used that experience has shown to be hazardous or unreliable.

(2) Further, the rule requires that the suitability of each questionable design detail and part must be established by tests.

b. Procedures.

(1) This rule is met partially by a review of service history of earlier model rotorcraft, or for a new model, review of service experience of models with similar design features. Specifically, this rule covers "features or details" such as the following:

Seat track-to-seat interface fittings. These fittings should have adequate locking devices to prevent both premature structural failure and premature unlatching.

(¡¡) Seat belt and harness should be of a type and construction that service experience has shown to be easy to don and unlatch and remove. They should also be of a type that is reliable, does not interfere with egress, and does not sustain unnecessary wear and tear under normal operations.

(iii) Metallic parts less than a certain thickness gauge and composite materials less than a certain number of plies should not be used. The minimum thickness and number of plies should be based to a large degree on service (normal wear and tear) experience with similar designs.

(2) The effects of service wear on the loading of critical components should be considered. Flight testing, ground testing, and analyses may be used in these considerations.

(3) Tests are required for details and parts which the applicant chooses to use after questions have arisen concerning their suitability.

242. § 27.603 (Amendment 27-16) MATERIALS.

a. Explanation. The rule requires that the suitability and durability of materials, the failure of which could adversely affect safety, must be determined by three-fold considerations:

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(1) Considerations based on experience or tests.

(2) By meeting approved specifications.

(3) By taking into account environmental conditions such as temperature and humidity.

b. Procedures.

(1) Where possible, materials that meet widely accepted specifications such as AISI, SAE, MIL, or AMS and alloys which have favorable experience or tests should be used. Where company developed materials are used, approved specifications are required to ensure the developed properties are duplicated in each lot of material.

(2) Environmental conditions may be taken into account by service experience, coupon testing, full-scale testing, or a combination of testing and experience, MIL-HDBK's -5, -17, and -23 include some environmental effects and contain reference to additional methods of testing for environmental effects.

(3) Section 27.613 concerns strength properties and design values. (See Paragraph 248 of this document.)

243. § 27.605 (Amendment 27-16) FABRICATION METHODS.

a. Explanation. The basic requirement of this rule is that the methods of fabrication must produce sound structure and produce it consistently.

(1) A process specification is required for fabrication processes requiring close control.

(2) A test program is explicitly required for each new aircraft fabrication method.

b. Procedures.

(1) The approved specifications required by this rule may either be established government/industry specifications such as MIL, AISI, ASIM, or SAE; or the specifications may be company-developed proprietary specifications. Sufficient data should be provided to the FAA/AUTHORITY aircraft engineering offices to show that the desired features are provided by the process specification. In addition, sufficient process controls, inspections, and tests should be coordinated with FAA/AUTHORITY manufacturing inspection personnel to ensure that continued quality of the process is provided.

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(2) In addition to the examples given by the rule; i.e., gluing, spot welding, and heat treating process, specifications should also be prepared for types of welding other than spot welding, for platings of metals, for protective finishes (other than decorative), for sealing, and for unique fabrication methods such as those used for composite materials.

(3) The required test programs should consider static strength effects, fatigue strength effects, and environmental effects as appropriate to the processes.

244. § 27.607 (Amendment 27-4) FASTENERS.

a. Explanation. Section 27.607 of Amendment 27-4 requires dual locking removable fasteners in critical locations. A nonfriction locking device is specifically required in any bolt subject to rotation, as stated in the rules.

b. Procedures. Advisory Circular 20-71 contains information, procedures, and means of complying with § 27.607 of Amendment 27-4.

245. § 27.609 PROTECTION OF STRUCTURE.

a. Explanation. The structure should be suitably protected as specified in the rule to maintain its design strength. Ventilation and drainage provisions must be provided as specified in the rule. Overboard drains should be furnished for corrosive or waste liquids. Drains for flammable fluids are specified in other rules such as §§ 27.999 and

27.1193.

b. Procedures.

(1) The structure may be preserved, painted, or treated with chemical films to protect it from strength deterioration. An approved process specification should be used for these types of treatments.

(2) Parts may be plated or chemically treated, such as anodized, for protection.

An evaluation and substantiation may be required to ensure the structure or parts are not adversely affected during, or as a result of, the plating or treatment process.

(§ 27.605 concerns approval of process specifications and fabrication methods.)

(3) Plating or material surface hardness or composition changes may require fatigue substantiation to ensure the fatigue strength is not altered or is otherwise properly assessed. An approved process specification should be used for these types of treatments.

(4) To prevent water accumulation, drain holes should be placed at possible dams such as bulkheads and at low points in the fuselage and in the stabilizing surfaces.

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(5) Control tubes and tubes used as primary mount structures (l.e., transmission support structure and engine mount structure) should be designed to prevent entry and collection of corrosive fluids or vapor, including water.

(1) A closed insert in each tube end may be used.

(il) A sealant applied around the tube ends and around each rivet head may be used.

(6) Overboard drains should discharge clear of the entire rotorcraft. Dyed water discharged in flight may be used to ensure fluids are properly drained.

(7) Drains or vents which handle corrosive fumes (such as battery case vent line) may incorporate a container with an agent to neutralize the fumes prior to venting overboard.

(8) Welded tubes should be flushed and sealed after welding in accordance with an approved process specification.

(9) Refer to AC 43-4, "Corrosion Control for Aircraft," for further procedures.

246. § 27,610 (Amendment 27-21) LIGHTNING PROTECTION.

a. Background. During the initial development and promulgation of the standards concerning the airworthiness of rotorcraft, it was not necessary to specify design features that would protect the rotorcraft from the meteorological phenomenon of lightning. This was due, in part, to the fact that rotorcraft were primarily operated in a VFR and nonicing environment. Also, a prudent pilot avoided thunderstorms where the possibility of encountering severe weather and a lightning strike was much greater. The construction, design, and operating environment of civil rotorcraft have changed markedly within the past two decades. Many rotorcraft are now authorized to fly IFR in all types of weather environment. One transport design has been approved for flight into known icing conditions. Additionally, many rotorcraft now use the same advanced technologies in structures and systems as do airplanes. Because of these facts, a specific rule on lightning protection of rotorcraft was adopted in Amendment 27-21. For further information, see the preamble of Amendment 27-21 (49 FR 44433; 11/6/84), Proposal 2-14. Section 27.610 is similar to § 25.581 which applies to the protection of structures of transport airplanes. However, the standard provides for specific protection of the aircraft structures as well as the systems of the rotorcraft. Section 27.610 is the standard referenced in the requirement for lightning protection of systems in § 27.1309(d) (see Paragraph 621 of this AC). In addition, the protection of fuel systems from the effects of lightning is found and referenced in Report DOT/FAA/CT-83/3, User Manual for AC 20-53A, Protection of Airplane Fuel Systems Against Fuel Vapor Ignition Due to Lightning, dated April 12, 1985.

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АС 27-1A 7/30/97

b. Explanation.

(1) The regulation requires that the rotorcraft must be protected against the catastrophic effects of lightning. This means that a lightning strike encounter should not prevent the continued safe operation of the rotorcraft.

(2) Paragraph 621 of this AC addresses the protection required for systems.

The protection of the rotorcraft structures may be accomplished in a similar fashion.

(3) The structural components of the rotorcraft should be designed in such a manner that the lightning current may be safely diverted or conducted through the rotorcraft without damaging any critical structure or without causing damage to noncritical structure, the failure of which would preclude the continued safe flight and landing of the rotorcraft. A radome or fin cap which explodes due to a lightning strike and results in catastrophic damage to main or tail rotors is a scenario of lightning damage to a noncritical structure which has catastrophic results.

(4) This type of strike effect on the rotorcraft is generally referred to as direct effects. Direct effects are damage which includes the burning, eroding, blasting, or structural deformation produced by the high currents of the lightning flash passing through the rotorcraft structure.

c. Procedures.

(1) Certification Plan. Although not a regulatory requirement, it is recommended that a formal written certification plan be used to assure regulatory compliance. The use of this plan is beneficial to both the applicant and the FAA/AUTHORITY because it identifies and defines an acceptable resolution to the critical issues early in the certification process. These are the usual steps to be followed when utilizing a certification plan:

(i> Prepare a certification plan which describes the analytical procedures and/or the qualification tests to be utilized to demonstrate protection effectiveness.

Test proposals should describe the rotorcraft and system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed test(s).

NOTE: The recommended reference for quantification of the lightning environment, the determination of the aircraft lightning strike zones, and the determination of appropriate test methods is SAE AE4L Committee Report, dated June 20, 1978, Lightning Test Waveforms and Techniques for Aerospace Vehicles and Hardware. Additionally, 304 Par 246

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information may also be found in the NASA publication No. RP-1008, Lightning Protection of Aircraft.

(ii) Obtain FAA/AUTHORITY concurrence that the certification plan is adequate.

Obtain FAA/AUTHORITY detail part conformity of the test articles and installation conformity of applicable portions of the test setup. Obtain FAA/AUTHORITY approval of the test proposal. A comprehensive test proposal may be used.

(iv) Schedule FAA/AUTHORITY witnessing of the test or tests proposed.

(v) Submit a test report describing all results and obtain FAA/AUTHORITY approval of each report prepared.

(2) Test Conditions. Refer to SAE AE4L Committee Report, dated June 20, 1978, and the NASA publication noted in Paragraph c(1)(i) to determine the appropriate test parameters.

(3) Aircraft Design Features and Criteria. MIL-B-5087B, Amendment 2 or later amendment, contains valuable information to assist the designer. Figure 6 in the specification contains fault current versus bond resistance information. Refer to the NASA publication noted above also.

Aluminum wire screen or mesh applied to the control or stabilizing surface and electrically bonded at each joint or juncture has been successful in conducting the current without serious damage.

(i¡) Metal skin surfaces combined with surface wire screen or mesh have been successful. Also, successful use of surface treatment has been reported. For composites, treatments such as the following have been used: flame spray coatings, aluminized glass, metal foil, metallized fabrics, and conductive paint.

(ili) Ball or roller bearings may be used to conduct the current at rotating joints. However, increased friction or possible seizure of the bearing may occur. The potential for this should be evaluated. Inspection and replacement criteria for possible damage should be addressed in the manual for continued airworthiness. Bearings are especially susceptible to pitting and internal arcing.

(iv) Report DOT/FAA/CT-86/8, April 1987, Determination of Electrical Properties of Grounding, Bonding, and Fastening Techniques for Composite Materials, may assist the applicant.

(4) Fuel Systems. Refer to Report DOT/FAA/CT-83/3 referenced in Paragraph 246a. For additional information on the lightning protection requirements for Par 246 305

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fuel systems for rotorcraft with a certification basis which includes Amendment 27-23 refer to Paragraph 449 of this AC.

247. § 27.611 INSPECTION PROVISIONS.

a. Explanation. The rotorcraft must have access panels or openings that will allow for proper maintenance and/or adjustment of the rotorcraft systems.

(1) The rule states: There must be means to allow close examination of each part that requires recurring inspection, adjustment for proper alignment and functioning, or lubrication."

(2) "Structural" or load-carrying access panels may be used to comply with the rule. Structural panels should have stencils or permanent labels (§ 27.1541(a)(2)) stating the panels must be installed prior to ground or flight operation.

(3) Holes or "nonstructural" access panels should be used whenever possible.

b. Procedures.

(1) The determination of compliance can be accomplished in conjunction with the following activities:

(i) Reviewing type design drawings.

(ii) Conformity inspections accomplished during certification testing.

(ili) Be evaluated during the control system proof and operation tests (§§ 27.681 and 27.683).

(iv) During type inspection tests and functioning and reliability testing.

(2) Equipment requiring frequent inspections (at less than 25-hour intervals), lubrication, or adjustments should be accessible through "nonstructural" doors. Areas or items requiring daily attention should be accessible through "nonstructural" doors since properly rated maintenance personnel are required to "open and close" or reinstall structural panels, and special design features, such as multiple pins and latches, are generally necessary for structural doors.

248. § 27.613 (Amendment 27-16) MATERIAL STRENGTH PROPERTIES AND DESIGN VALUES.

a. Explanation. The rule requires the use of materials that have a known minimum strength value. The structure must not be understrength and must be designed to minimize fatigue failure.

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(1) Material design values in certain specified documents may be used. The FAA/AUTHORITY may approve other material design values thus allowing the applicant greater flexibility in selection of materials by proving their strength properties and design values as stated in § 27.613(d).

(2) Other materials that may be new or are not included in the specified documents may be tested and design values established as provided by § 27.613(a) and (d).

(3) Section 27.613(d) requires the selection of materials that will retain design values and properties in the type of service environment and for the length of service time intended for the structure.

(4) Section 27.613(c) is an objective rule concerning minimizing fatigue failures and § 27.571 concerns quantitative fatigue substantiation requirements.

b. Procedures.

(1) The properties and design values in the documents noted in the rule may be used.

(2) MIL-HDBK-5, Metallic Materials and Elements for Flight Vehicle Structure, Chapter 9, contains procedures for establishing design values of additional materials.

Uniform means of presenting the data are also contained in this chapter.

(3) Design values and properties must include effects of the service environment and service time. An example is exposure at elevated temperatures on the ultimate tensile strength of 7079-T6 aluminum alloys as found in Figure 3.7.4.1.1(c) of MIL-HDBK-5.

(4) The probability of disastrous fatigue failures must be minimized. This may be accomplished by using design features usually identified as fail-safe features, such as the following, which were obtained from Advisory Circular 20-95.

Selection of materials with stress levels to provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks (lightly loaded structures).

(її) Use of multipath construction and the provision of crack stoppers to limit the growth of cracks.

(iii) Use of composite (multielement) duplicate structures so that a fatigue crack or failure occurring in one element of the composite (multielement) member will Par 248 307

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be confined to that element and the remaining structure will still possess adequate load-carrying ability.

(iv) Use of backup structure wherein one member carries all the load, with a second member available and capable of assuming the extra load if the primary member fails.

(v) Design to permit detection of cracks including the use of crack detection systems, in all critical structural elements before the cracks can become dangerous or result in appreciable strength loss, and to permit replacement or repair.

(5) Acceptable standards for pressurized containers or cylinders, such as cylinders of nitrogen, used to inflate emergency floats may be found in 49 CFR 178, Subpart C, §§ 178.36 through 178.68. Specifically, § 178.44 concerns standards for steel cylinders used in aircraft that are subjected to at least 900 PSI service pressure.

This standard includes strength, test, material property, inspection, quality, design features, identification, and inspection report requirements. As an example, § 178.44-14, entitled "Hydrostatic Test," requires that each cylinder must be (proof) tested to at least 5/3 times the service pressure. Section 178.44-16, entitled "Burst Test," also states that one cylinder taken at random out of each lot of cylinders shall be hydrostatically tested to destruction.

(6) Other design criteria may be developed and approved under the provisions of FAR Part 27 as a unique part of the aircraft type design.

248A. § 27.613 (Amendment 27-26) MATERIAL STRENGTH PROPERTIES AND DESIGN VALUES.

a. Explanation. Amendment 27-26 added explicit probability standards criteria to § 27.613(b). This amendment also provided for testing or proving the strength of selected individual items rather than conducting coupon tests to develop generic material strength properties that would be used for design purposes.

b. Procedures. The basic procedures of Paragraph 248 of this AC still apply, except:

(1) Probability criteria common with MIL-HDBK-5D are explicitly allowed to determine strengths for metallic materials whose data are not available in MIL-HDBK-5D. These specific probability criteria should be used in conjunction with MIL-HDBK-17B whenever determining material strength properties for non-metallics.

(Also, reference Paragraph 788 of this AC).

(2) New § 27.613(e) provides for the premium selection of materials. The premium selection of materials method uses a specimen from each individual item (part) to determine its properties before its use is allowed. This is a highly specialized 308 Par 248

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and possibly costly method which applies only to parts that have areas available from which specimens can be obtained without destroying the part. The rotorcraft type design data of those parts made from premium selection should have the necessary information, such as a minimum allowable strength, on the drawing.

249. § 27,619 SPECIAL FACTORS.

a. Explanation.

(1) This is a general rule to complement other rules. Special factors are employed for reasons cited in the rule to ensure an airworthy aircraft structure. The

1.5 ultimate load factor in § 27.303 is multiplied by a special factor as specified in the rule.

(2) Specific factors are prescribed for castings and fittings in §§ 27.621 and

27.625, respectively. Factors may be prescribed for bearings with free clearance as stated in § 27.623. In addition, any other factor may be prescribed "to ensure that the probability of the part being understrength because of the uncertainties specified in § 27.619(a) is extremely remote."

b. Procedures.

(1) One example of fitting factor use follows:

1,000-pound limit design load × 1.15 fitting factor x 1.5 ultimate load factor equals 1,725-pound ultimate design load.

(2) Other specific factors may be similarly applied. Refer to §§ 27.623 and

27.625.

(3) Other factors may be imposed as cited in the rule. Advisory Circular 20-107, Paragraphs 5 and 6, are examples of requiring tests of component and subcomponent structure to account for variability of strength and stiffness of composite structures. Factors appropriate for the particular design are obtained and used in substantiation of the composite structure.

(4) The rule complements §$ 27.603 and 27.613. Regardless of the rule invoked, the variability of the material and/or assembly properties should be accounted for.

250. § 27.621 CASTING FACTORS.

a. Explanation. Casting design, test, and inspection criteria are included in this rule for critical and noncritical structural castings. Hydraulic or other fluid containers are Par 248A 309

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АС 27-1A 7/30/97 not subjected to "structural loads" but are subject to pressure testing as a part of hydraulic or other flight systems. Critical and noncritical castings are defined in the rule.

(1) Factors, tests, and inspections are specified for structural castings.

Additional factors, tests, and inspections may be applied, as prescribed by § 27.603, § 27.605, or § 27.613, for foundry quality control.

(2) For castings that have surfaces subject to bearing structural design loads, the casting factor need not exceed 1.25 with respect to bearing stresses and need not be used with respect to the bearing surfaces if the bearing factor of § 27.623 exceeds the applicable casting factor.

(3) Critical castings must have a casting factor not less than 1.25 and must receive 100 percent inspection as specified including radiographic inspection. Static test requirements are also specified in addition to the inspection requirements.

(4) Noncritical structural castings may have a casting factor as small as 1.0 with attendant increased inspection and quality control requirements. Use of larger casting factors reduces the inspection and quality control requirements.

(5) Structural static and fatigue substantiation, by test or analysis, is still required in addition to any casting static tests required by this rule.

b. Procedures.

(1) The rotorcraft castings should be classified as critical or noncritical or nonstructural or fluid container as soon as possible in the certification program. The applicant should then be prepared to propose the tests required for certification.

(2) The casting factors and associated inspection requirements dictated by § 27.621(c) and (d) are shown in the following chart:

310 Par 250

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**INSPECTION REQUIREMENTS**

CRITICAL CASTINGS NONCRITICAL CASTINGS <(2)> <(3)>

**CASTING**

FACTOR FAA REQUIRE- OTHER FAA REQUIRE- OTHER RANGE MENT 27.621(c) CLASSIFICATION MENT 27.621(d) CLASSIFICATION <(1)>

2.01 OR <(7)> <(4)>

**GREATER**

1.50 TO <(7)> <(5)>

2.00

1.250 TO <(7)> <(8)> <(6)>

1.499

1.00 TO NOT NOT <(7)> <(8)>

1.249 ALLOWED ALLOWED <(9)> <(1)> Ultimate load = Casting factor x 1.5 x limit load. CAUTION: For casting factor range of 1.25 to 1.5 see yield test requirements of NOTE 8.

The mechanical properties to be used for analysis shall be based on the tabulated values of MIL-HDBK-5 or other approved sources, ref. § 27.613.

<(2)> Critical castings are those castings whose failure would preclude continued safe flight and landing or result in injury to any occupant, ref.

§ 27.621(c).

<(3)> Noncritical castings are castings other than those defined by NOTE 2.

<(4)> Each casting shall receive 100 percent visual inspection.

<(5)> Each casting shall receive 100 percent visual and reduced magnetic particle or penetrant inspection or approved equivalent methods.

Par 250 311

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<(6)> Each casting shall receive 100 percent visual and reduced radiographic and magnetic particle or penetrant inspection or approved equivalent methods.

<(7)> Each casting shall receive 100 percent inspection by visual, radiographic, and magnetic particle or penetrant inspections or approved equivalent methods.

<(8)> Three sample castings shall be static tested and shown to meet:

No failure at 1.25 x 1.5 x limit load, and no yielding at 1.15 x limit load.

<(9)> Castings shall be procured to a specification that guarantees the mechanical properties of the material in the casting and provides demonstration of these properties by test of coupons cut from the castings on a sampling basis.

This chart may be included in the casting test proposal report. It is recommended that the applicant include in the test proposal report additional information such as shown in Paragraph 250b(3).

(3) The casting test report may include the following sections or items in a Part I of the report. The report may also have a Part II that contains the test results as shown in the following example report. The following sections are a recommended format content of the report. Appropriate changes should be made as desired to accommodate the applicant's system.

**EXAMPLE OF REPORT**

INTRODUCTION This report presents the proposal for the static test of the castings used on the Model XYZ. The castings will be tested in compliance with Federal Aviation Regulations § 27.621. The purpose of this test is to substantiate the structural strength of the castings used on the Model XYZ. Part Il of this report, which will be published after static tests have been completed, will present test results.

All test specimens will be selected as radiographic standards of acceptance for the particular castings (see Test Specimen). Additional information on selecting the specific castings may be included in the test specimen section of this report.

Load sheets giving direction and magnitude of loads for each of the castings are presented in numerical order by part number at the end of this report. The test loads 312 Par 250

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7130/97 AC 27-1A and design criteria for the castings are discussed in detail in the test loads section of this report.

The test loads will be applied and reacted using mating aircraft parts or special fixtures which simulate the mating parts. The methods and apparatus to be used for the static tests of the castings are discussed in the apparatus and method section of this report.

Testing will be conducted in... (location).

TEST SPECIMEN The castings which will be tested are listed in numerical order in Table I. Those castings which, after structural analysis, show less than a 1.5 casting factor will be tested. All directions are given with reference to a forward facing position in the rotorcraft.

On the basis of a radiographic examination, the three castings which are of the poorest acceptable quality in the first production lot of castings will be selected as test specimens. The poorest of the three castings will be selected as the initial test casting and its radiograph or ASTM standard will be used as the standard for accepting future castings of the particular part unless later standards are approved. Three castings must be tested for each critical condition for each part.

Conformity Inspection Each machined casting will be subjected to an FAA/AUTHORITY conformity inspection prior to testing to determine compliance with the type design drawings. A conformity report for each casting may be incorporated in Part II, Test Results, of this report.

The test specimen will be permanently marked or defaced after testing to preclude its use on a rotorcraft.

See Table I for an example of a convenient means of listing castings.

TEST LOAD The test load(s) to be applied to each casting represents the critical loading condition(s) for that casting. The critical conditions on each of the castings were determined by the design criteria and substantiating data approved by the FAA/AUTHORITY.

The design criteria for all of the castings to be static tested may fall into one of two categories. The load factors and structural acceptability requirements for each category are discussed below. Casting factors that are included on the load sheets of each part do not apply in the discussion below. (See Paragraph 250b(2) for casting factors.)

Par 250 313

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Castings Designed to Limit Load Conditions A structural analysis of each test casting showing the critical design limit load conditions is given in the data (reference report number here). The load factors for the static test of the castings are as follows:

1.15 x design limit load = design yield load

1.50 x design limit load = design ultimate load Castings Designed Only to Crash Landing Conditions The castings in this category were designed using a crash landing load factor for the design ultimate load. The design yield load criteria of 1.15 x limit load need not apply to these castings. The test loads for these castings may be given in terms of design ultimate load on the individual casting load sheets shown in Part I of this report.

Test Procedures Depending on the results of the initial static test of each casting, the following procedure will be used.

a. If in the initial test of critical castings the casting is found to have a casting factor of 1.5 (1.5 x design ultimate load), the casting will be considered acceptable and no further tests will be conducted.

b. If in the initial tests) the critical casting is found to have a casting factor less than 1.5 but equal to or greater than 1.25, two additional castings will be tested for each critical load condition. Each must also show a minimum casting factor of 1.25.

c. If in the initial test, or in one of two additional tests, a casting shows a casting factor less than 1.25 times design ultimate or yields prior to reaching 1.15 times design limit load, the casting will be redesigned and retested. The yield criteria are also applicable to the first two procedures with the exception of critical castings designed to crash landing conditions.

TEST APPARATUS AND METHOD The Model XYZ casting static tests will be conducted using fixtures designed to simulate the installation of the castings in the aircraft. Where practical, mating aircraft parts will be used to apply and react test loads. When practical, the static tests will be conducted with mating castings assembled when the critical loads for the mating castings are compatible; otherwise, fixtures simulating the mating parts may be designed and fabricated for the tests. Assembly hardware used to mount test castings will be the same as hardware used on the rotorcraft. All bolt torques and other assembly notes will conform to the type design assembly instructions.

314 Par 250

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7/30/97 АС 27-1A The tests will be conducted using calibrated load measuring devices such as hydraulic cylinders and pressure gages, load cells, strain gage bridges, or dead weights.

Deflections of the casting may be measured using graduated dial indicators or scales in all tests. The deflection indicators will be based or mounted on the casting and will measure casting deflection only when possible; otherwise, the indicators will be based on the fixture and measure deflection of the casting relative to the fixture. Deflection readings will be made at 20 percent increments of limit load through 100 percent of limit load and at 115 percent of limit load. These increments may be changed if necessary.

Permanent deformation readings will be made after relieving 115 percent and 150 percent of limit load.

See Figure 250-1 as an example of a load sheet.

Par 250 315

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P +5,000 # TENS.

ULTIMATE LOADS -1,000 # COMP Tension Comp.

Tension Reaction Compression Reaction FIGURE 250-1

**EXAMPLE OF CASTING LOAD SHEET**

RETRACT ACTUATOR SUPPORT - LANDING GEAR 316 Par 250

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Include spherical bearing with clamped-up bolt and a link in the test setup to confirm the stability. Loads are based on a jam condition with actuator operating at 1,700 PSI pressure maximum.

A 1.25 casting factor is included in these loads.

These loads were derived from data in approved structural loads and analysis report.

**END OF SAMPLE REPORT**

(4) The format of the previous guidance material may be changed to accommodate the applicant's method of data presentation.

(5) Nonstructural castings may be tested and included in the test report.

(6) Cast fluid containers, including hydraulic fluid containers, may be tested as prescribed in other rules of FAR Part 27 and a test proposal and test results report may be included in the casting test report, or an appropriate report may be referenced for convenience. We recommend use of one report to contain test data or reference to test data for all castings used on the rotorcraft.

Par 250 317

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**TABLEI EXAMPLE**

**CASTINGS TO BE STATIC TESTED FOR MODEL XYZ REF. LOAD**

**MACHINE OR SHEET CASTING NO. ASSY. NO. NAME AND LOCATION MATERIAL FIG. NO.**

Base Assembly, Pilot's Collective Column 318 Par 250

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251. § 27.623 BEARING FACTORS.

a. Explanation.

(1) The rule requires use of a minimum bearing factor in free fit joints to account for effects of typical relative motion. A minimum value is not specified in the rule. The factor, appropriate for the application, is applied to the ultimate bearing strength of the softest material used as a bearing. A definition of free fit (clearance fit) is noted in Paragraph 251b(7) below.

(2) A bearing factor, appropriate for the application, shall be used unless a larger factor is used.

(3) For reference, specific bearing factors are contained in § 29.685(e) for transport rotorcraft control system joints subject to angular rotation. These factors are applied to the ultimate bearing strength of the softest material used as a bearing in the control system. Control systems ball, roller, or needle bearings are covered by § 29.685(f) for transport rotorcraft.

(4) MIL-HDBK-5D, Paragraph 8.3, refers to design standards for plan or journal bearings or bushings. These standards are found in Air Force Systems Command Design Handbook AFSC DH-2-1, Airframe, Chapters 3 and 6.

b. Procedures.

(1) Control system joint bearings are discussed under Paragraph 284, § 27.685, of this AC, but the bearing factors are noted here for convenience. For transport rotorcraft control systems, § 29.685(e) requires a 2.0 bearing factor for cable systems and a 3.33 bearing factor for push-pull systems other than ball and roller bearing systems. The manufacturer's static, non-Brinell rating of ball and roller bearings should not be exceeded. Use of this for normal category rotorcraft is recommended.

(2) A landing gear pivot, grease-lubricated, plain bearing is one example of a free fit subject to pounding or vibration. A bearing factor of 2.0 may be used or another factor may be proven for a grease-lubricated plain bearing or bushing to account for the anticipated higher loads caused by pounding or vibration. See Paragraph 251b(6) for recommendations on ball or roller bearings.

(3) A typical engine mount bolt installation with a plain bearing having a free or loose fit (not interference fit) is another example of a sleeve bearing application subject to a design bearing factor. As an EXAMPLE OR ILLUSTRATION, a bearing factor of

1.85 may be applied to the design loads on the softest material reacting the bearing loads. A different but appropriate factor will be acceptable. The design limit load may be calculated for the example of a 0.312-inch-diameter bolt in a 2-inch-long bearing.

Par 251 319

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The bearing projected area is 0.312 × 2 = 0.624-square-inches. The design limit load is 3,000 pounds. The design limit bearing stress is 3,000 pounds/0.624-square-inch x

1.85 = 8,894 PSI. If a free or loose fit is not used; i.e., tighter than free fit, a bearing factor is not required. See Paragraph 251a(4) for bearing factors.

(4) Military standard part specifications, MS 21240, "Bearing, Sleeve Plain, TFE Lined," and MS 21241, "Flanged Bearing, Sleeve Plain, TFE Lined," contain allowable load ratings, static and dynamic, that apply to the particular use of the bearing. An appropriate bearing factor should be applied to the static rating. Military Specification (MIL-B-8943A, Amendment 3, "Bearing, Sleeve, Plain, and Flanged, TFE Lined" (temperature range -65° F to +250° F) shows that MS 21240 and MS 21241 sleeve bearings have been superseded by MS 1934/1 and MS 81934/2 sleeve bearings, respectively. Military Specification MIL-B-81934, Amendment 2, , "Bearings, Sleeve, Plain and Flanged, Self-Lubricating," uses TFE liners. These bearings are intended for use in a temperature range from -65° F to +325° F. Whenever a sleeve bearing is used, an appropriate bearing factor should be applied to the static rating that is contained in the specification or standard. Other sleeve bearings are contained in standards NAS 72 through NAS 77, NAS 537, and NAS 538. The installation design information is only contained in standards NAS 72 through NAS 74. These types of plain sleeve bearings are designed for clamping to the shaft or bolt with relative motion occurring on the bearing outside diameter. An appropriate bearing factor is required for the application.

(5) The minimum fitting factor 1.15, specified by § 27.625, must be applied as specified to account for load distribution at the fitting. This fitting factor need not apply to plain or journal "bearings" whose "bearing factor" exceeds 1.15.

(6) For airframe and landing gear structural joints, the manufacturer's static, non-Brinell rating of ball and roller bearings should not be exceeded. ABEC Class 1 bearings or better quality bearings may be used in airframe structural joints and landing gear; ABEC Class 3, 5, or 7 bearings should be used in rotor pivot joints. The non-Brinell rating includes consideration of the bearing factor, and no other bearing factor is necessary.

(7) A free fit was described in American Standards Association (ASA) Standard B4a-1925. The "free fit" clearances and tolerances of this old standard are now called Class RC6, Medium Running Fit, in ASA Standard B4.1, 1955. As an illustration using these standards, a 1-inch diameter shaft and a plain sleeve bearing would have a clearance ranging from 0.0014 to 0.0040 inch.

252. § 27.625 FITTING FACTORS.

a. Explanation. A 1.15 factor is specified to ensure that the calculated load and stress distribution within any fitting is conservative. Application of the factor is excluded or is an exception as stated in the rule.

320 Par 251

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b. Procedures.

(1) The factor may be applied to the calculated load or stress for the fitting.

(2) The structural design substantiating data should include the fitting factor and where applicable should include, but not be limited to, the rotor system. The rotor system includes the rotor blade attachments, rotor head and hubs, and boosted control system elements. Other typical areas that may be considered are tail rotor gearbox attachment, tailboom to fuselage fittings, transmission pylon attachments, and landing gear attachment to the rotorcraft.

(3) The fitting factor is not required in the following applications:

(i) Joints such as continuous joints in metal plating, welded joints and scarf joints in wood.

(її) Elements proven by limit or ultimate load tests such as nonboosted control system parts.

(iii) Elements for which a larger load factor is used such as a casting factor, a 1.33 retention factor when required for seats and safety belts, a fatigue factor, bearing factor or special factor greater than 1.15, crash load factors that are the only design case, and crash load factors that exceed limit load factors x 1.5 x 1.15.

(iv) Elements for which the failure mode does not affect safety of flight or occupant satety.

253. § 27.629 FLUTTER.

a. Explanation. The rule requires that the rotorcraft "be free from flutter under each appropriate speed and power conditioni

b. Procedures. Freedom from flutter is to be shown for the entire rotorcraft with special attention to the blades, fins, and stabilizers.

(1) Flutter is defined as an aeroelastic instability resulting primarily from coupling of flap and pitch bending modes.

(2) Freedom from flutter may be shown by analysis or by appropriately instrumented flight flutter tests.

(3) The flight load survey proposal submitted for compliance with § 27.571 may also contain tests to fulfill compliance with § 27.629.

Par 252 321

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АС 27-1A 7/30/97

(4) Flight loads survey data or flight flutter test data should be reviewed to ensure that excessive oscillatory deflections of rotors or surtaces will not be encountered.

(5) Sensitivity analyses should be conducted to ensure that normal wear in the pitch change mechanisms of the main rotor blades and tail rotor blades does not reduce the effective stiffnesses sufficiently to cause flutter.

322 Par 253

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**254.-264. RESERVED.**

Intentionally Left Blank 323 (thru 333) Par 254

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**SECTION 16. ROTORS**

265. § 27.653 (Amendment 27-2) PRESSURE VENTING AND DRAINAGE OF ROTOR BLADES.

a. Explanation. The rule requires each rotor blade to be provided with venting and drainage means (i.e., holes, etc.) or else the blade must be sealed and designed to withstand internal pressure.

b. Procedures. Although the rule provides for venting and drainage features, recently certificated blades have been designed to be sealed and to sustain the "maximum pressure differentials expected in service." For modern blade designs, the internal pressure buildup due to environmental effects and centrifugal acceleration effects (near the tip) can be readily sustained with moisture sealing accomplished. The use of sealed blades is highly advantageous and recommended because of the possibility for severe corrosion damage resulting from trapped moisture and because of the difficulty in finding internal corrosion damage by use of field level inspections.

266. § 27.659 (Amendment 27-2) MASS BALANCE.

a. Explanation. The rule requires that mass balancing of rotors and blades be provided, as necessary, to prevent excessive vibration and flutter. Further, the rule requires structural substantiation of the mass balance installation.

b. Procedures.

(1) The weight, geometry, and location of rotor and blade mass balance devices are determined as the requirements of §$ 27.571 and 27.629 are met.

(2) The structural substantiation should show static strength to meet the maneuver and gust loads of §§ 27.337, 27.339, and 27.341. In addition, the main rotor loads of § 27.547(c) should be substantiated. The fatigue strength of the mass balance devices (including structural supports) should meet the requirements of § 27.571.

(3) In addition to the appropriate strength requirements, some recent designs have included features which trap the balance weight inside a limited area even if the primary attachment means (adhesive, bolts, etc.) fail. This type of design feature is recommended because of the severe loading environment to which balance devices are subjected.

267. § 27.661 (Amendment 27-2) ROTOR BLADE CLEARANCE.

a. Explanation.

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(1) The rotors, main and tail, must not strike other parts of the rotorcraft during any operating condition.

(2) Section 27.411 concerns protection of the tail rotor from a ground strike.

b. Procedures.

(1) The applicant should have drawings or sketches of the rotorcraft that show an adequate minimum clearance between the rotors, main and tail, and parts of the rotorcraft. Probable flexing of the rotors should be considered in determining the minimum clearance.

(2) During parts of the FAA/AUTHORITY-conducted flight test program, frangible devices (wood dowels) or other means of measuring clearance, may be requested to confirm that the clearance shown in the drawings or sketches is adequate in certain operating conditions. Balsa wood dowels may be clamped to the aft part of the fuselage within the rotor arc. If the devices are intact after autorotation landing tests and other tests involving typical abrupt, cyclic, and rudder pedal displacements, the clearance should be satisfactory and compliance obtained. If such measuring devices are used, the type inspection report should contain a record of clearance found during the tests. It is not necessary to precisely determine the clearance but only necessary to determine "enough clearance" as stated in the rule.

268. § 27,663 (Amendment 27-2) GROUND RESONANCE PREVENTION MEANS

a. Explanation.

(1) This rule, adopted in Amendment 27-2 and revised in Amendment 27-26 requires reliability and damping action investigation for the ground resonance prevention means. The probable range of variations in service, not just the allowable range, must be established and investigated as prescribed. This probable range includes operation on the ground, and other appropriate landing surfaces applicable to the rotorcraft design shall be considered. Quantitative test data are generally obtained in compliance with this rule, but analysis or tests may be used.

(2) Appropriate maintenance information should be included in the maintenance manual (also called instructions for continued airworthiness).

(3) Paragraph 99 of this document concerns demonstrating freedom from ground resonance during certain applicant and TIA verification evaluations or tests of the rotorcraft. Section 27.241 complements the requirements of § 27.663. As noted in Paragraph 99 of this document, a specific requirement for a ground vibration survey was removed from CAR Part 6. However Section 27.663 was adopted by Amendment 27-2 to investigate possible sources of ground resonance and to assure Par 267 335

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the reliability of the ground resonance prevention means, i.e., dampers, if necessary, to preclude occurrence of ground resonance. The total rotorcraft system is evaluated under this rule.

(4) Viscous dampers have been used for many years to prevent ground resonance. Modern rotorcraft designs may also use elastomeric dampers and may use elastomeric bearings in the rotor head and rotor pylon attachment to the airframe. The rule also requires investigation of the probable range of variations of these dampers, whether viscous or elastomeric, and these bearings to preclude ground resonance.

(5) Ground resonance can occur due to flexibility in the rotor pylon restraint system as well as with landing gear flexibilities. See Paragraph b(2) for an explanation.

An analysis may be done to show the effect of the rotor pylon mount stiffness on ground resonance stability. If the analysis shows that rotor pylon mount stiffness could affect ground resonance, the evaluation should include variations in stiffness and damping of the rotor pylon restraints that may occur in service (reference "Ground Vibrations of Helicopters," M.L. Deutsch, JAS, Vol. 13, No. 5, May 1946).

b. Procedures.

(1) The reliability of the means for preventing ground resonance may be substantiated as stated in the rule. An analysis report or a test proposal and subsequent test report may be used to show compliance. The probable ranges of damping restriction are an important part of the assessment. The test may be conducted in conjunction with the testing required by § 27.241. See Paragraph 99.

(i) Analysis and tests may be used.

(її) Reliable service history of identical or closely similar systems may be used. The materials and fluids used, clearance or fits, seals, and physical installation are important items to be evaluated and considered for "closely similar" systems.

(ili) Testing of the complete rotorcraft may be used to prove that malfunction of a single means or member of the damping system will not cause ground resonance. One method of demonstrating acceptable compliance is by removing all or most of the fluid from a damper and considering the allowable ranges of damping of the other parts of the rotorcraft damping system while operating the rotorcraft throughout the rotor speed range from start to maximum rotor speed. Investigation of elastomeric dampers may require innovative test procedures and preliminary discussions of these prior to preparation of a test proposal. The rotorcraft cyclic control should be displaced as noted in Paragraph 99 of this document to assure that the possible rotorcraft resonance frequencies are excited. If vibrations are damped in all tests, the damping system is satisfactory. Each critical rotor damper and landing gear damper must simulate a malfunction to comply with the rule. The testing discussed, however, could be come very extensive if one were to attempt to test all combinations of all 336 Par 268

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maintenance adjustments of all components which contribute to the prevention of ground resonance, while at the same time rendering each of the pertinent components ineffective in turn and then repeating all of the maintenance tolerance testing each time.

Fortunately, rational analytical methods are available which will permit the evaluation of such combinations so that only the combinations with the least amount of margin used are physically tested.

(2) The pylon damper variation can affect ground resonance. The variations in stiffness and/or damping of pylon mounts should be evaluated except the pylon mounts on contemporary conventional rotorcraft may have little influence on "classical" ground resonance stability. The dynamics of the rotorcraft on its landing gear is generally established by the airframe properties and the landing gear properties under the influence of the rotor system, with the "pylon" having little or no effect. For air or flight resonance, the rotor generally couples with the rigid body modes of the fuselage. For a specific design, a relatively simple analysis may be used to show the effect of the pylon mount system stiffness on air and ground resonance stability, and if not important, variations in the system may be omitted from the test program.

(3) The probable ranges of damping must be established and investigated as prescribed and noted in Paragraph 268(b). An approved test proposal and test results report should be used for complying with § 27.663(b). If wheel landing gear is used on the rotorcraft, the probable ranges of tire pressure or the lowest probable tire pressure should be stated in the test proposal and effects of the tire pressure investigated during the test. See Paragraph 99, § 27.241, concerning tests and instrumentation of the test associated with complying with § 27.241. The instrumentation noted in Paragraph 99 also applies to § 27.663(b).

(4) If the wheel landing gear is equipped with wheel brakes, the evaluation should include brakes "on" and "off." The nose or tail wheel should be locked and unlocked if it swivels to evaluate any possible adverse effects of this feature.

(5) Any maintenance procedures should be included in the "recommended" part of that manual. See Appendix A, FAR Part 27.

268A. § 27.663 (Amendment 27-26) GROUND RESONANCE PREVENTION MEANS.

a. Explanation. Amendment 27-26 clarifies that analysis as well as tests may be used to show freedom from ground resonance after malfunction or failure of a single means of ground resonance prevention. This amendment primarily clarifies that the probable range of damping should be established as well as investigated.

b. Procedures. The procedures of Paragraph 268 of this AC continue to apply with the addition of the need to document the establishment of the probable range of damping of ground resonance prevention means. Acceptable tire and oleo minimum Par 268 337

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АС 27-1A 7/30/97 and maximum pressures as well as other identified factors should be documented in maintenance instructions if necessary to maintain the desired characteristics.

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**269.-277. RESERVED.**

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**SECTION 17. CONTROL SYSTEMS**

278. § 27.671 GENERAL.

a. Explanation.

(1) The rule requires basically that controls operate easily and smoothly and provide positive response of the rotorcraft from control input.

(2) In addition, the rule requires that incorrect assembly be prevented by special design features or special markings.

b. Procedures.

(1) Easy, smooth operations of controls are substantiated by the operations tests of § 27.683 and the FAA/AUTHORITY flight testing under TIA procedures.

Positive response of the rotorcraft to control inputs is also evaluated during company flight testing and FAA/AUTHORITY TIA flight testing to the requirements of §§ 27.141 through 27.175.

(2) To meet the requirement that incorrect assembly be prevented, the preferred method is providing design features which make incorrect assembly impossible. Typical design features which can be used are different lug thicknesses, different member lengths, or significantly different configurations for each system component. In the event that incorrect assembly is physically possible (because of other considerations), the rule may be met by the use of permanent, obvious, and simple markings. Permanent (durable) decals or stencils may be used.

(3) Design features of the control systems are checked when reviewing the type design drawings. During the proof and operation tests of §§ 27.681 and 27.683, the controls should be thoroughly reviewed for possible incorrect assembly and for any required markings supplied for compliance with this standard.

279. § 27.672 (Amendment 27-21) STABILITY AUGMENTATION.

**AUTOMATIC, AND POWER-OPERATED SYSTEMS.**

a. Explanation.

(1) This rule requires that the pilot be made aware of stability augmentation, automatic, or power-operated system failures which could lead to an unsafe condition.

Examples of clearly distinguishable warnings include, but are not limited to, an obvious aircraft attitude change following the failure or an audio warning tone. A visual indication itself may not be adequate since detection of a visual warning would normally require special pilot attention. The use of devices such as stick pushers or shakers is 348 Par 278

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not acceptable as a warning means. However, this rule is not intended to eliminate the use of such devices for other purposes. Examples of automatic control systems other than a stability augmentation system would be a pitch axis actuator used for the purpose of demonstrating compliance with longitudinal static stability requirements or a fly-by-wire elevator. For control systems where a series actuator malfunction could degrade control authority, a means should be provided to the pilot to determine actuator alignment (see § 27.1329(b)).

(2) The corrective flight control input following a system failure should be in the logical direction. For example, a malfunction resulting in a nosedown pitch of the aircraft should require a corrective cyclic control input in the aft direction. The system deactivating means does not have to be located on the primary flight control grips;

however, it should be easily accessible to the pilot. Malfunctions and subsequent recoveries must be shown throughout the operating envelope of the aircraft. In a case where control authority is decreased following a malfunction, a reasonable flight envelope must be defined wherein compliance with controllability and maneuverability requirements can be demonstrated. This reduced flight envelope must be presented in the flight manual. Compliance with trim and stability characteristics is not required following a malfunction; however, a pilot workload assessment should be made to show that a mission can be safely continued to completion following the worst case single failure.

b. Procedures. A discussion of malfunction test procedures is presented in Paragraph 775b(6). Controllability and maneuverability test procedures are addressed in Paragraph 81.

280. § 27.673 (Amendment 27-21) PRIMARY FLIGHT CONTROL.

a. Explanation. This regulation basically defines primary flight controls as "those used by the pilot for immediate control of pitch, roll, yaw, and vertical motion of the rotorcraft." This regulation was generated to clarify the application of § 27.1555 which requires markings for controls other thaniprimary flight controls or controls) whose function is obvious."

b. Procedures. The primary flight controls; i.e., cyclic stick, collective, and tail rotor pitch control pedals are excluded from the marking requirements of § 27.1555.

281. § 27.674 (Amendment 27-26) INTERCONNECTED CONTROLS.

a. Explanation. A new § 27.674 is added by Amendment 27-26 which requires that the rotorcraft be capable of safe flight and landing after a malfunction, failure, or jam of any auxiliary interconnected control.

b. Procedures.

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(1) Section 27.674 requires that the rotorcraft be shown to be capable of safe flight and landing after a malfunction, failure, or jam of an auxiliary control interconnected with a primary control. The section does not apply to interconnected primary controls; e.g., cyclic and collective controls.

(2) Examples of auxiliary controls covered by this section may include certain autopilot or stability augmentation or trim system components. Section 27.1309 methods may be used in determining failure effects of autopilot and stability augmentation "system" components. For components whose purposes are solely mechanical functions, the procedures associated with § 27.571 for components such as the main rotor may be used.

(3) If an engine control could jam and result in a collective control jam, the controls should be designed to relieve that connection.

282. § 27.675 (Amendment 27-16) STOPS.

a. Explanation.

(1) Stops are required to prevent unrestrained movements of pilot/autopilot inputs from causing interferences or overloads.

(2) The rule requires that the stop must be located to not appreciably affect the control system range of travel due to wear, slackness, or takeup adjustments.

(3) Each stop is required to withstand loads corresponding to design conditions.

(4) In addition, each main rotor blade, if appropriate for the design, must have stops to limit its travel about its hinge points. For rotors with hingeless design, stops may be provided as appropriate to limit blade travel. Loads which result from the blade hitting the stops (during starting or stopping the rotor or during any large but allowable pilot control inputs such as autorotation cyclic flares or when subjected to ground gusts, etc.) shall not overload the stops nor any rotor component.

b. Procedures.

(1) Stops are generally provided in the cockpit area and near any controllable surface end of the control system (i.e., main rotor hub, tail rotor hub, and stabilizer activators). For systems with control coupling or series actuators, stops have been located farther downstream (away from the cockpit) to permit increased control output during malfunction (hardover) or extreme control position cases.

(2) Location of stops in close proximity to each end of a control system will allow the stop to provide its function most efficiently without undue deflections between 350 Par 281

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the stop and its adjacent surface or its adjacent cockpit control lever or pedals. The location of stops close to the control lever or surface will help meet the requirement that the stop (and its function) not be appreciably affected by wear, slackness, or takeup adjustments. Consideration should be given to limiting the total amount of takeup adjustments of both the stop and the control systems to preclude a hazardous adjustment of the control surface range of travel by either normal or extreme takeup adjustment.

(3) Each stop is to be substantiated for critical design conditions from either pilot effort, aerodynamic loads, hydraulic loads, and other critical loads, as applicable.

The stops can be substantiated for limit loads by the tests of § 27.681.

(4) The stops to limit the main rotor blade about its hinge points should be positioned to prevent the blades from striking any part of the structure, particularly during startup and shutdown operations. These stops should also limit the flapping of the static main rotor blades of the rotorcraft when they are subjected to ground gusts and rotor wash from nearby taxiing rotorcraft. Provisions should be made to prevent overloading the stops or the blade under conditions of ground gusts and rotor wash effects or during autorotational landing flares. The need for provisions to prevent possible overloads due to ground gusts and close taxiing by adjacent rotorcraft and by autorotational landing can be determined using the instrumented flight load survey aircraft by hover-taxiing another rotorcraft near the instrumented aircraft and by conducting autorotational landing flares with the instrumented aircraft. Substantiation for the final main rotor flapping stop design can be demonstrated by similar tests.

(5) If features of design are added to the main rotor stop assembly which activate certain portions of the stop assembly only on the ground to meet the requirement that the blade not hit the droop stop during any operation other than starting and stopping the rotor, such features of design must be substantiated to reliably operate by both ground tests and flight tests, as appropriate. Wear and rigging tolerances should be considered in these demonstration tests.

283. § 27.679 CONTROL SYSTEM LOCKS.

a. Explanation. The rule requires that if control system locks are provided, means are necessary to prevent the rotorcraft from taking off with the locks engaged or, once airborne, to prevent the locks from engaging in flight.

b. Procedures. Two main procedures may be used to meet the requirements of this rule.

(1) The first procedure is to provide a means to disengage the lock "automatically" as the pilot operates the controls. If this method is used, the means must disengage the lock in a manner that it will not automatically re-engage during flight under normal pilot operations. The means may be physical removal of the locking Par 282 351

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device from close proximity to the control system interface with deliberate crew action necessary to return the device to the control system interface, or the means may be that the mechanism geometry and/or actions prevent locks from engaging in flight.

(2) The second procedure which may be used is to provide locks which so limit rotorcraft operations that it is impossible to take off with the locks engaged. Acceptable means are features which prevent engine startup or which restrict collective control operations to prevent sufficient lift for takeoff.

284. § 27.681 LIMIT LOAD STATIC TESTS.

a. Explanation.

(1) The rule requires static tests of the control system in showing compliance with limit load requirements.

(2) The tests are specified to include each fitting, pulley, and bracket of the control system being tested and to include the "most severe loading."

ference $ 27.623) be shown by individual tests orby analyses for control system joints subject to motion.

b. Procedures.

(1) Compliance with the requirements of this rule is obtained by static tests conducted on either a static test airframe or on a prototype flying ship. In either case, conformity of the control system and related airframe is necessary to validate the tests.

(2) The rotor blades or aerodynamic surfaces may be used to react pilot effort loads through the control system, or they may be replaced with fixtures. If fixtures are used, they should be evaluated for geometric and stiffness efforts to ensure test validity.

(3) The loads to be applied during the limit load static tests are specified in §§ 27.395, 27.397, and 27.399. The loads are applicable to collective, cyclic, yaw, and rotor blade control systems as well as any other flight control systems provided by the design.

(4) Although Part 27 does not explicitly specify the bearing factors to be used in control system rotating joint tests or analyses, the factors of § 29.685 have been used in past programs. These factors are 3.33 for push-pull systems and 2.0 for cable systems for joints with plain bearings and manufacturers' ratings for ball and roller bearings.

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285. § 27.683 OPERATION TESTS.

a. Explanation. The rule requires that the control system be free from jamming, excessive friction, and excessive deflection. An operational test is required in which specified loads are applied at the pilot controls and carried through an operating control system.

b. Procedures.

(1) Compliance with the requirements of this rule is obtained by use of a test setup similar to that used for the limit load tests of § 27.681, except the load reactions at the blades (or surfaces) must allow for movement of the blades (or surfaces) as the system is operated through its operating range.

(2) Fixtures are normally affixed to the surfaces (or replace the surfaces) to allow pulley arrangements which provide for movement under load. These fixtures should be evaluated to ensure that system loads up to limit will be applied during the full range of operations of each system.

(3) Each flight control system should be operated through its entire range under a light load and under limit load. As the controls are being operated, the system should be checked for jamming, excessive friction, and excessive deflection. Excessive deflection includes deflection sufficient to contact other systems or structures. Also (in agreement with CAM 04.331/04.43.11), FAA/AUTHORITY policy has been to consider excessive the deflection of a control system under limit load which exceeds approximately one-half of the system travel from neutral to the extreme stop. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system.

286. § 27.685 (Amendment 27-11) CONTROL SYSTEM DETAILS.

a. Explanation. The rule requires that the control system be designed to prevent chafing, jamming, and interference from cargo, passengers, loose objects, or the freezing of moisture. Specifically, means are required in the cockpit to prevent the entry of foreign objects into places where they would jam the system, and means are required to prevent the slapping of cables or tubes against other parts.

b. Procedures.

(1) The geometry of the control system components and their installations are the primary control to prevent chafing, jamming, and interference. The control system from cockpit to surface should be checked for clearances both unloaded and loaded.

The control system should be checked under load during both the limit load static tests (reference § 27.681) and the operational tests of § 27.683. Location of guides or fairleads and pulleys may be used in cable systems to prevent chafing and interference Par 285 353

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AC 27-1A 7130/97 with other structure. Generally, tubes should clear adjacent structure by location and design geometrical considerations. If supplemental means are provided to assure the tubes do not chafe or interfere, the means should be evaluated for possible jamming.

(2) Rubber (or other elastomeric) boots connected to both the cockpit control arm or shalt and to the floor are acceptable means to prevent the entry of foreign objects into underfloor areas where they may cause jamming of controls. Control systems should, in general, be routed around cargo compartments. If routing of the control system components is in or near cargo areas, the control system components should be protected by bulkheads, panels, or other enclosures which have sufficient strength and stiffness to prevent possible interference with the control system components when subjected to cargo loading and handling deflections.

(3) Control system details should be reviewed for possible moisture collection.

Areas should drain free. Exposed or open control areas should drain tree and areas ot possible freezing moisture collection should not accumulate ice that would cause a jam of the controls. Simulated or actual ice collection on the controls may be used to prove questionable features. The areas to be considered for moisture collection include both external and internal areas where moisture may accumulate by direct impingement of water, entrapment of water particles, or condensation of moisture.

286A. § 27.685 (Amendment 27-26) CONTROL SYSTEM DETAILS.

a. Explanation. Amendment 27-26 adds §§ 27.685(d), (e), and (f) for cable systems, control system joints, and bearings, which are compatible with the same pre-existing paragraphs of § 29.685 except cables of 3/32 inch diameter are allowed by this section rather than the minimum of 1/8 inch diameter required by § 29.685 for transport rotorcraft.

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

(1) The latest revisions of MIL-HDBK-5D do not explicitly give approved pulley-cable combinations, but appropriate MIL specifications are referenced in Chapter 8.3 of MIL-HDBK-5D for use in determining pulley-cable combinations and ratings.

(2) Adhere to the ratings, factors, and alignment as specified.

(3) Provide inspection means as specified for the control system.

(4) Close fitting pulley guards are required for cable systems.

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287. § 27.687 SPRING DEVICES.

a. Explanation.

(1) This standard for control systems ensures that springs and spring devices used to prevent flutter, control oscillations, or vibrations are either -- (i> Reliable (failure is extremely remote); or

(ii) The failure is not critical to the rotorcraft.

(2) Tests simulating service conditions are required in either instance.

b. Procedures.

(1) Springs and spring devices used in the control system, including balance springs, should be identified early in the certification program.

(2) Whenever a spring cannot be proven by observation or analysis that it is "not critical," then ground or flight tests may be required.

(3) Springs that are critical to safe operation may be subject to fatigue substantiation to prove they are reliable for the operating conditions imposed in service.

(4) Springs used in conjunction with hydraulic actuator spool valves may be subject to the standards of § 27.695.

288. § 27.691 AUTOROTATION CONTROL MECHANISM.

a. Explanation

(1) Rotorcraft designs generally have a main rotor blade collective pitch control system that does not have detents or other devices to limit pitch control in the control midrange. Autogyro and other rotorcraft designs may include detents or other finite position control for collective pitch control. This rule requires that the control design allows rapid entry into autorotation after a power failure.

(2) Section 27.33 contains standards concerning establishment and control of the main rotor speed limits. The standard requires flight tests and demonstrations. The standard also concerns rotorcraft design features that are related to control of the main rotor speed limits.

(3) Other design requirements for control systems are contained in § 27.685.

b. Procedures.

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(1) If high and low main rotor pitch stops are employed in the collective control and if the control may be rapidly moved from one limit to the other, compliance is shown.

(2) If detents or intermediate stops are employed, the pilot must be able to easily and readily override, disconnect, remove, or bypass the device to allow rapid autorotational entry prior to exceeding transient low speed rotor limits. An early assessment of the design may be accomplished by the flight test personnel with the evaluation completed in the Type Inspection Authorization (TIA) test program.

(3) It is acknowledged that modern rotorcraft designs may have an autorotational VNE that is lower than "power-on" VNE or normal cruise speed. For rotorcraft designs with this characteristic, the speed must be reduced after entry into autorotation. No relief from the rule is required since many phases of operation occur at speeds less than power-on VNE. For example, a critical phase of flight occurs during takeoff. Rapid entry into autorotation is essential during this phase also.

(4) The features of the autorotational control mechanism and ability to control the rotor speed within the design limits for any rotorcraft will be evaluated as an integral part of the TIA test program.

289. § 27.695 POWER BOOST AND POWER-OPERATED CONTROL SYSTEM.

a. Reference Regulations. The following sections of Part 27 are either incorporated in the provisions of § 27.695 or are otherwise applicable to power boost and power-operated control systems:

(1) Section 27.307 Proof of structure.

(2) Section 27.571 Fatigue evaluation of fight structure.

(3) Section 27.671 Control system.

(4) Section 27.681 Limit load static tests.

(5) Section 27.687 Spring devices.

(6) Section 27.685 Control system details.

(7) Section 27.861 Fire protection of structure controls and other parts.

(8) Section 27.863 Flammable fluid fire protection.

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(9) Section 27.1301 Function and installation.

(10) Section 27.1309 Equipment, systems, and installations.

b. Explanation.

(1) The rule requires an alternate system if a power boost or power-operated control system is used.

(2) The alternate system must, in the event of any single failure in the power portion of the system, or in the event of failure of all engines:

(i) Be immediately available.

(ii) Allow continued safe flight and landing.

(3) The alternate system may be:

(i) A duplicate power portion of the system; or

(il) A manually operated mechanical system.

(4) The power portion of the system includes:

The power source (such as hydraulic pumps); and

(il) Items such as valves, lines, and actuator.

(5) The failure of mechanical parts (such as piston rods and links) must be considered unless their failure is extremely improbable.

(6) The jamming of power cylinders must be considered unless their jamming is considered extremely improbable.

c. Procedures. It is assumed in the following discussion that the power boost or power-operated control system being utilized is a typical aircraft hydraulic system.

(1) The rule requires, without respect to the probability of failure, an alternate system for the power portion of the system. The power portion of the system, by example in the rule, includes hydraulic pumps, valves, lines, and actuators. It has also been interpreted to include seals, servo valves, and fittings.

(2) If a duplicate power portion of the system is used to meet the requirements of the rule, the requirements may be met by providing a dual independent hydraulic system, including the reservoirs, hydraulic pumps, regulators, connecting tubing, hoses, Par 289 357

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servo valves, servo-valve cylinder, and power actuator housings. There must be no commonality in fluid-carrying components. A break in one system should not result in fluid loss in the remaining system.

(3) Dual actuators should be designed to ensure that any single failure in the duplicated portion of the system, such as a cracked housing, broken interconnecting input, or broken interconnecting output link, does not result in loss of total hydraulic system function.

(4) A manually operated mechanical system may be used as the alternate system to a single hydraulic system if, after the loss of the single hydraulic system, the pilot can control the rotorcraft without undue mental or physical fatigue in any normal maneuver for a period of time as long as that required to effect a safe landing.

(5) The substantiation of the various system components should include consideration for operation in the normal and alternate system modes.

(6) The "extremely improbable" criteria noted in § 27.695(c) for failure of mechanical parts may be satisfied by performing component fatigue testing and establishing a service life through this technique.

(7) Fatigue substantiation of the control actuator is required under § 27.571 and should consider both the stresses imposed by flight loads and the stresses imposed by hydraulic pump pressure pulses. Flight loads factored in a conservative way may be an acceptable means to take into account both effects.

(8) The possibility of jamming of the power cylinder may be shown as "extremely remote" through a failure analysis that considers every possible system component failure such as, but not limited to, ruptured lines, pump failure, regulator failure, ruptured seals, clogged filters, jammed servo valves, broken interconnecting servo valve inputs, broken interconnecting output links, etc.

(9) Three acceptable means to meet the requirements of § 27.695(a)(2) could be as follows:

(i) Provide two transmission-driven hydraulic pumps, provided the pumps are driven by the transmission during all flight conditions including autorotation.

(ii) Use two electrically-driven hydraulic pumps if electrical power is available to drive the pumps with all engines failed. If this approach is used, the battery must be capable of running both pumps plus all other required equipment necessary for continued sate flight.

(iii) Use a single transmission-driven pump and an electrically driven pump.

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7/30/97 АС 27-1A

**290.-296. RESERVED.**

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**297. RESERVED.**

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7/30/97 АС 27-1A SECTION 18. LANDING GEAR

298. § 27.723 SHOCK ABSORPTION TESTS.

a. Explanation.

(1) Limit and "reserve energy" drop tests are required as prescribed in §§ 27.725 and 27.727, respectively. These tests may be conducted on the complete rotorcraft or on units consisting of wheel, tire, and shock absorber in their proper relation. For rotorcraft with skid landing gear, the tests may be conducted on the complete rotorcraft or on a simulated fuselage with the complete skid landing gear system.

(2) The rotorcraft must be designed to limit load factors that equal or exceed the limit load factor substantiated by these drop tests. In practical application, the rotorcraft may be designed to a limit load factor, such as 2.8g. Thus, it is necessary that the limit landing load factor derived from the landing gear drop tests be equal to or less than 2.8g. If not, the rotorcraft must be redesigned for the higher load factor derived from the drop tests. It must be shown in accordance with § 27.723 that the limit load factors selected for design under § 27.473 will not be exceeded in landings with the limit descent velocity corresponding to the drop height specified in that section. In addition, reserve energy absorption capacity of the landing gear must be shown for a descent velocity of 1.22 times the limit descent velocity selected under § 27.473 by increasing the drop height to 1.5 times the "limit" drop height. The test requirements or procedures outlined in Part 27 for obtaining the landing load factors are empirical;

however, these procedures are based on and supported by satisfactory experience.

(3) As stated in § 27.725(c), each landing gear unit should be tested in the attitude simulating the landing condition that is most critical from the standpoint of the energy to be absorbed by it. For wheel landing gear designs, the level landing or tail down landing and level landing with drag are generally the most critical attitude. A test of more than one attitude may be required to comply with the standard.

(4) Drop tests are required. If analytical methods and/or means are proposed by the applicant, the data presented for approval must be equal to or conservative with respect to that data obtained from physical drop tests. Section 21.21(b)(1) concerns "equivalency" determinations. Presenting an acceptable means of "equivalency" here would circumvent the necessary scrutiny of an analytical method or means and is also beyond the scope of this document.

b. Procedures. The test plan or proposal must be approved prior to official FAA/AUTHORITY tests unless satisfactory resolution of outstanding proposal or conformity inspection items can be accomplished after the test.

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(1) The following headings would be a typical table of contents for the test proposal, and a generalized explanation of the contents that may be included under each of these headings for a wheel landing gear follows.

Purpose. The regulations to which compliance is being shown by the drop tests should be identified (usually §§ 27.723, 27.725, and 27.727). Also, the rotorcraft landing gear, including the wheels and tires to be dropped, should be positively identified in the report by the manufacturer's or applicant's previously FAA/AUTHORITY-approved drawing, technical standard orders (TSO's), or other identifying FAA/AUTHORITY-approved data as applicable.

(її) Description of test setup. This section should present a description of the test fuselage or jig, method of attaching landing gear to jig, and type of accelerometer to be used to measure load factors. Proof of calibration of accelerometer should be available. The accelerometer should be mounted at the aircraft CG if a free drop of the aircraft is used or as close as practical to the centerline of the main shock absorbing component of each landing gear (oleo, strut, etc.) if each gear is tested separately. The description of the test jig, including platforms on which the gears are to be dropped, should be defined by sketches in addition to the required mathematical calculations. This data should show that the landing gear will be at the proper attitude, relative to the platform, on impact for the particular landing condition.

Drawings or other approved data from which the geometry is taken should be referenced in the proposal. The tire and oleo pressures at the time of the test should be specified. The method of measuring the deflection of the tire plus the vertical travel of the axle under impact should be described. This measurement may be accomplished by telescoping tubes attached to the point on the jig that would measure the total (tire and oleo) vertical deflection of the landing gear. Other vertical and horizontal deflections should be measured as required to determine if the landing gear has experienced permanent deformation after each drop test. The effect of surface roughness should be considered. Smooth surfaces tend to give maximum deflections where rough surfaces tend to restrict deflection and to result in maximum values of Nz.

Preliminary company drop tests (at less than limit drop height) may be used to determine the critical surtace roughness, or engineering evaluations may be used (without tests) when the gear configurations are such that the critical surface condition can be analytically determined (or when the load factor is shown to be negligibly affected by surface roughness). NACA Report 1154, dated 1953, contains information that surface coefficients of friction may vary from 0.4 to 0.7. Skid landing gear standards, § 27.501(c), indicate an acceptable coefficient of friction is 0.5. A wheel landing gear design standard, § 27.479(b), indicates an acceptable coefficient of friction is 0.25. In the case of a small rotorcraft, the entire aircraft may be dropped. This may be accomplished by establishing pivot points at the main gear axles for the tail (or a point forward of the nose gear) drops and a pivot point at the tail (or nose gear) axle for the main gear drops. It is the responsibility of the applicant to distribute the aircraft inertia items, including added weight to get the proper effective drop weight (We) at the 368 Par 298

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landing gear, so that no local failures of the aircraft occur as a result of the limit or reserve energy drop tests.

(iii) Test data. Computations for the required drop height (h) and the effective drop weight (We) should be shown for each design level landing and tail down landing condition in compliance with §$ 27.479 and 27.481. The computations should be in accordance with § 27.725(a) for h and § 27.725(b) for We for the limit drop tests.

We and h are computed in accordance with § 27.725 for the limit drop test and with § 27.727 for reserve energy drop test. The computation of the static weight on the gear being dropped (W, WN, or WT) and used in the computation of We should be shown.

This static weight is defined as WM, WT, or Wy for the main gears, tail gear, or nose gear, respectively, in § 27.725(d). It should be shown that the critical CG and proposed certificated maximum landing weight have been used in the computation of Ww, WT, or WN. The computation of the slope of the platforms required for the inclined reaction conditions should be presented also.

(iv) Test results. The results of the test are based on the values of We, h, d, W, and L used and obtained for each drop test and the value of N; obtained from the accelerometer. These results should be summarized, and the method of computing the aircraft limit inertia load factor should be shown for each drop in accordance with § 27.725(d). A print or copy of the film or other recording trace from the accelerometer, if not a direct readout type of accelerometer, should be included in the test results.

Each critical condition should have several preliminary drops, as many times as required, to obtain reasonable correlation.

(2) Skid landing gear may be tested using similar procedures except a level landing attitude drop test is all that is required by § 27.501. The design load conditions specified in § 27.501(c) through (t) are derived from this level drop test condition.

Section 27.501(a)(2) and (3), contain special considerations for skid landing gear.

(її) Section 27.501(a)(2) specifies that structural yielding of elastic spring members under limit load is acceptable. This yielding or deformation is a means of absorbing the landing impact. For skid landing gear that uses oleo or other types of shock absorbers, the standard does not allow structural yielding under limit load.

During the limit load and reserve energy (ultimate for skid landing gear with elastic spring numbers) drops, the yielding energy absorbing members will probably deform or yield. After a limit drop test, the gear may be used for a reserve energy drop at the discretion of the applicant, but a gear that has been subjected to a reserve energy drop should not be used unless it can be shown that no yielding has occurred in that gear.

(3) Wheel landing gear is tested in attitudes prescribed in Paragraph 298a(3).

Each unit, nose or main gear, is generally tested separately.

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(4) Skid landing gear is tested in attitudes prescribed in Paragraph 298a(3).

Due to the construction of skid landing gear, the complete skid landing gear is tested as a unit. Thus, the level landing with drag condition is probably the critical attitude for the forward cross-tube and its attachments. The level landing condition is probably the critical attitude for the aft cross-tube and its attachments.

(5) An FAA/AUTHORITY or FAA/AUTHORITY designated or delegated person need only witness the drop tests for "record" or "compliance." Preliminary or developmental drops do not require an FAA/AUTHORITY witness.

299. § 27.725 LIMIT DROP TEST.

a. Explanation. Limit drop tests in the critical aircraft attitude or critical attitude of each gear are required for the landing gear. The drop height must be at least 8 inches, which equates to a 393-foot-per-minute (free fall) vertical descent speed. Rotor lift may be simulated, and an effective mass may be used in the drop test as prescribed.

b. Procedures. See Paragraph 298, § 27.723, of this advisory circular.

300. § 27.727 RESERVE ENERGY ABSORPTION DROP TEST

a. Explanation.

(1) In addition to the limit drop tests, a reserve energy drop test is required.

The landing gear must not collapse in this test to the extent that the fuselage impacts the ground. Fracture (to separation) of landing gear parts is considered collapse of the landing gear. This test is not an ultimate load drop test for the landing gear, except as specified in § 27.501(a)(3) for certain skid landing gear designs using elastic spring members.

(2) All other types of landing gear must be substantiated for design ultimate loads in addition to this reserve energy drop test.

(3) Shock absorbing devices, such as oleos, must not "bottom" during the reserve energy drop test. "Bottoming" occurs when displacement of the device no longer occurs with increasing load.

(4) Requirements for proof of the landing gear and airframe structure are found in §§ 27.305, 27.307, and 27.473.

b. Procedures. See Paragraph 298, § 27.723, of this advisory circular.

370 Par 298

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