Rotorcraft Flight Manual

CSP−902RFM207E−1

� RFM Revision900 Series

Transmittal/( 2 blank)

Manual: CSP−902RFM207E−1, Rotorcraft Flight ManualModels: MD900 (902 Configuration with PW 207E) HelicopterOriginal Issue: 28 July 2000Reissue 1: 2 December 2008

The attached manual is a reissue of the MD902 (with PWC 207E engines) flight manual. Re-fer to the ‘‘Summary of Revisions’’ page for revised material included in this reissue.

FILING INSTRUCTIONS:

Remove all pages in Sections I thru XI (do not remove STCsupplements or weight and balance records in Section VI) includingfront matter, and replace them with the new flight manual.

Operators with questions regarding the new manual may contact:Mike Zale at 480−346−6558 from 6:00 AM to 2:30 PM MST.


Model MD900 (902 Configuration with PW 207E)

F90−001

R O T O R C R A F T F L I G H T M A N U A L

REGISTRATION NO: _________________SERIAL NO: _________________

Cover


Copyright � 1999−2008 by MD Helicopters, Inc. All rights reserved under the copyright laws.

F−iOriginal

FAA APPROVED

ROTORCRAFT

FLIGHT MANUAL

for

Model MD900

THE FAA APPROVED ROTORCRAFT FLIGHT MANUAL CONSISTS OF THE FOLLOWING SECTIONS.

SECTION II − LIMITATIONSSECTION III − EMERGENCY PROCEDURESSECTION IV − NORMAL PROCEDURESSECTION V − PERFORMANCE DATASECTION X − OPTIONAL EQUIPMENTSECTION XI − CATEGORY−A OPERATIONS

The helicopter must be operated in compliance with the operating limitations as set forth in section II ofthis manual and any additional limitations from Section X as a result of an installed optional equipmentitem.

Sections III, IV, V, X and XI contain recommended procedures and data and are FAA approved.

THIS MANUAL MUST BE KEPT IN THE HELICOPTER AT ALL TIMES.

Title Page

(902 Configuration with PW 207E)

Type Certificate No. H19NM

Approved By:_____________________________________

Manager, Flight Test Branch, ANM−160L

Federal Aviation Administration

Los Angeles Aircraft Certification Office

Transport Airplane Directorate

Original Approval Date: July 28, 2000

Reissue #1 ________________________

CSP−902RFM207E−1 ROTORCRAFT FLIGHT MANUAL

MD900 (902 Configuration with PW 207E)

Original F−iiReissue 1

LOG OF REVISIONS BY DATEFAA / NON−FAA REVISIONS

REVISION NUMBER AND DATE

Original Issue 28 July 2000.

Revision 1 30 May 2001. . . . . .

Revision 2 02 November 2001. . . . . .

Revision 3 26 December 2002. . . . . .

Revision 4 13 February 2003. . . . . .

Revision 5 23 April 2003. . . . . .

Revision 6 22 September 2003. . . . . .

Revision 7 13 January 2004. . . . . .

Revision 8 8 July 2005. . . . . .

Revision 9 21 August 2007. . . . . .

Revision 10 20 February 2008. . . . .

Revision 11 11 April 2008. . . . .

Reissue #1 2 December 2008. . . . .

Approved By:

_______________________

Manager, Flight Test Branch, ANM−160LFederal Aviation AdministrationLos Angeles Aircraft Certification OfficeTransport Airplane Directorate

CSP−902RFM207E−1ROTORCRAFT FLIGHT MANUAL


Original F−iii/( F−iv blank)Reissue 1

APPROVING AUTHORITIES

Joint Aviation Authorities (JAA)

This manual was approved by the JAA.

European Aviation Safety Authority (EASA)

The MD900 and this manual were accepted by EASA based on the JAA approvallisted above. Subsequent revisions of this manual are approved by EASA whoissues a four−digit approval number. See MDHI web page http://www.mdhelicop-ters.com, publications link for EASA approval number and instructions.

National Agency of Civil Aviation (Brazil)

This Aircraft Flight Manual is approved by the FAA on behalf of the NationalAgency of Civil Aviation for Brazilian registered aircraft, in accordance withthe Regulamentos Brasileiros de Homologação Aeronáutica" (RBHA) 21, Section21.29.

/ (Initial FAA Approval Signature/Date)Manager, Flight Test Branch, ANM−160LFederal Aviation AdministrationLos Angeles Aircraft Certification OfficeTransport Airplane Directorate

CSP−902RFM207E−1ROTORCRAFT FLIGHT MANUALMD900 (902 Configuration with PW 207E)

F−vOriginal

Reissue 1

TABLE OF CONTENTS

PARAGRAPH PAGECover 1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Title Page F−i. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Log of Revisions By Date F−ii. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Approving Authorities F−iii. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Summary of Revisions to the Rotorcraft Flight Manual F−xi. . . . . . . . . . . . . . . . . . . . . . . .

List of Effective Pages F−xii. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section I − General1−1. Introduction 1−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−2. Scope 1−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−3. Organization 1−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−4. Method of Presentation 1−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−5. Definition of Terms 1−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−6. Abbreviations 1−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−7. Multi−Purpose Utility Operations 1−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−8. Technical Publications 1−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−9. Rotorcraft Certification 1−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−10. Pilot’s Briefing 1−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−11. Dimensions 1−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−12. Conversion Charts 1−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section II − Limitations2−1. Flight Restrictions 2−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−2. Environmental Operating Conditions 2−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−3. Airspeed Limitations 2−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−4. Weight Limitations 2−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−5. Center of Gravity (CG Envelope) 2−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−6. Rotor Brake Limitations 2−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−7. Rotor Speed Limitations 2−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−8. Transmission Limitations 2−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−9. Power Plant Limitations 2−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−10. Generator Limitations 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

ROTORCRAFT FLIGHT MANUALMD900 (902 Configuration with PW 207E)


F−vi OriginalReissue 1

PARAGRAPH PAGE2−11. Starter limitations 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−12. Fuel System Limitations 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−13. Integrated Instrumentation Display System (IIDS) 2−10. . . . . . . . . . . . . . . . . . . . . . .

2−14. Decals and Placards 2−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section III − Emergency and Malfunction Procedures3−1. General 3−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−2. Caution and Warning Annunciators and Audio Tones 3−2. . . . . . . . . . . . . . . . . . . . . . .

3−3. Engine Emergencies 3−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−4. Emergency Landing Procedures 3−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−5. EEC Malfunctions 3−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−6. Engine Starting − Manual 3−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−7. Engine/Aircraft Shutdown − Manual 3−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−8. Fire Emergencies 3−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−9. Flight Control Malfunctions 3−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−10. Pitot/Static System Malfunction: Single or Dual Pitot Tube Installation 3−23. . . . .

3−11. Engine and Generator Malfunction Indications 3−24. . . . . . . . . . . . . . . . . . . . . . . . . . .

3−12. Transmission Malfunction Indications 3−28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−13. Fuel System Display Advisories 3−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−14. Caution and Warning Advisories 3−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−15. Other Malfunction/Advisories 3−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−16. Vibrations 3−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−17. Emergency Egress 3−42. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−18. Emergency Equipment 3−43. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section IV − Normal Procedures4−1. Preflight Requirements 4−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−2. Pilot’s Daily Preflight Check 4−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−3. Pilot’s Preflight Check 4−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−4. Engine Pre−Start Cockpit Check 4−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−5. Engine Starting − Automatic 4−24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−6. Engine Runup 4−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−7. Before Takeoff 4−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−8. Normal Takeoff 4−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


F−viiOriginal

Reissue 1

PARAGRAPH PAGE4−9. Cruise 4−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−10. Slow Flight/Approach 4−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−11. Landing 4−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−12. Engine/Aircraft Shutdown − Normal 4−28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−13. Post Flight 4−30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−14. Noise Impact Reduction Procedures 4−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−15. Flight With Doors Removed or Cabin Doors Open 4−32. . . . . . . . . . . . . . . . . . . . . . . . .

4−16. One Engine Inoperative Training 4−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−17. Fuel System 4−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section V − Performance Data5−1. General 5−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−2. Noise Characteristics 5−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−3. Density Altitude Chart 5−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−4. Airspeed Calibration 5−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−5. Best Rate of Climb Speed 5−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−6. Rate of Climb and Descent − OEI 5−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−7. Rate of Climb − AEO 5−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−8. Hover Ceiling, AEO 5−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−9. Hover Ceiling, OEI 5−50. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−10. Height Velocity Diagram 5−53. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−11. Power Assurance Check − Automatic 5−54. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−12. Power Assurance Check − Manual 5−56. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section VI − Weight and Balance Data6−1. Weight and Balance Characteristics 6−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−2. Load Limits and Balance Criteria 6−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−3. Equipment Removal or Installation 6−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−4. Longitudinal Weight and Balance Determination: Passenger Configuration 6−8. .

6−5. Longitudinal Loading of Cargo 6−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−6. Permissible Lateral Loadings − Passenger Configuration 6−10. . . . . . . . . . . . . . . . . . .

6−7. Lateral Loading of Cargo 6−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−8. Internal Loading of Cargo 6−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .



F−viii OriginalReissue 1

PARAGRAPH PAGE

Section VII − Systems Description7−1. Helicopter Exterior Description 7−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−2. Fuselage 7−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−3. Tailboom and Empennage 7−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−4. Landing Gear 7−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−5. Main Rotor System 7−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−6. Flight Controls 7−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−7. Hydraulic Systems 7−20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−8. Propulsion System 7−24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−9. Engine Air Intake and Inlet Particle Separator (IPS) 7−28. . . . . . . . . . . . . . . . . . . . . . .

7−10. Engine Power Management System 7−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−11. Fuel System 7−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−12. Fire Extinguishing System 7−34. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−13. Electrical System 7−36. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−14. Environmental Control 7−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


7−16. IIDS Data Storage 7−49. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−17. Balance Monitoring System 7−52. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−18. IIDS Menu Structures 7−53. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Section VIII − Handling, Servicing, and Maintenance8−1. Hoisting, Lifting, and Jacking 8−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−2. Towing and Moving 8−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−3. Parking and Storage 8−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−4. Access and Inspection Provisions 8−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−5. Servicing 8−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−6. Aircraft Cleaning 8−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−7. Cockpit Door Removal 8−34. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−8. Cabin Seats: Removal/Installation 8−36. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−9. Copilot Flight controls 8−37. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−10. Engine Charts 8−38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−11. Special Operational Checks and Procedures 8−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


F−ixOriginal

Reissue 1

PARAGRAPH PAGE

Section IX − Additional Operations and Performance Data9−1. Abbreviated Checklists 9−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

9−2. Fuel Flow vs Airspeed 9−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

9−3. International Civil Aviation Organization (ICAO) Noise Levels 9−26. . . . . . . . . . . . . .

Section X − Optional Equipment10−1. General Information 10−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−2. Listing − Optional Equipment 10−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−3. Compatibility − Combined Optional Equipment 10−2. . . . . . . . . . . . . . . . . . . . . . . . . . .

10−4. Optional Equipment Performance Data 10−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−5. Operating Instructions: Air Conditioning (P/N 900P7250302−103) 10−3. . . . . . . . . .

10−6. Operating Instructions: Controllable Landing/Search Light 10−7. . . . . . . . . . . . . . . .

10−7. Operating Instructions: Rotorcraft Cargo Hook Kit 10−13. . . . . . . . . . . . . . . . . . . . . . . .

10−8. Operating Instructions: Windscreen Wipers 10−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−9. Operating Instructions: Supplemental Fuel System 10−25. . . . . . . . . . . . . . . . . . . . . . .

10−10. Operating Instructions: Rescue Hoist 10−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−11. Operating Instructions: Removable CoPilot Controls 10−49. . . . . . . . . . . . . . . . . .

10−12. Operating Instructions: Airframe Fuel Filter 10−55. . . . . . . . . . . . . . . . . . . . . . . . .

10−13. Operating Instructions: SX−16 Night Sun with Aft Mount 10−59. . . . . . . . . . . . .

10−14. Operating Instructions: RDR−1400C Weather Radar 10−65. . . . . . . . . . . . . . . . . .

10−15. Operating Instructions: LEO−II−A5 Observation System 10−71. . . . . . . . . . . . . .

10−16. Operating Instructions: Annunciator panel 10−75. . . . . . . . . . . . . . . . . . . . . . . . . . .

10−17. Operating Instructions: Moving Map Navigation Systems 10−77. . . . . . . . . . . . .

10−18. Operating Instructions: W.E.S.T Battery Protection System 10−81. . . . . . . . . . . .

10−19. Operating Instructions: SX−16 Night Sun with EPMS Mount and Laser Pointer 10−83. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−20. Operating Instructions: Smoke Detector 10−89. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−21. Operating Instructions: Crew Door Modification with Quick Release Mechanism 10−93. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .



F−x OriginalReissue 1

PARAGRAPH PAGE

Section XI − Category A OperationsPart I General 11−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−1.1. General 11−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−1.2. Definitions − Category A Takeoff 11−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−1.3. Definitions − Category A Landing 11−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Part II Limitations 11−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−2.1. Clear Airfield, Heliport and Elevated Helipad 11−5. . . . . . . . . . . . . . . . . . . . . . . .

11−2.2. Maximum Takeoff and Landing Weight Limits 11−6. . . . . . . . . . . . . . . . . . . . . . . .

Part III Takeoff and Landing Procedures 11−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−3.1. Clear Airfield Takeoff Procedures 11−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−3.2. Heliport/Elevated Helipad Takeoff Procedures 11−13. . . . . . . . . . . . . . . . . . . . . . . .

11−3.3. Landing Procedures − Clear Airfield, Heliport and Elevated Helipad 11−15. . .

11−3.4. Equipment Malfunctions 11−18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Part V Performance Data 11−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.1. Takeoff Performance 11−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.2. Takeoff Distance Required 11−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.3. Continued Takeoff FLight Path 11−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.4. Landing Performance − Open Airfield 11−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.5. Landing Performance − Heliport/Elevated Helipad 11−27. . . . . . . . . . . . . . . . . . . .

Part IX Additional Operations 11−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−9.1. Category A OEI Training 11−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .



Original F−xiReissue 1

SUMMARY OF REVISIONSTO THE ROTORCRAFT FLIGHT MANUAL

NOTE: Revisions are listed below by number with appropriate remarks.

Section II pages marked [C]* indicate FAA approved color pages.Black−and−white reproductions of color pages are not considered to be “FAAApproved”.

REVISIONNUMBER REMARKS

Reissue 1 All pages now indicate the current reissue number above the effectiverevision number.

Sections I thru IV and VII thru XI: Added metric equivalents to En-glish measurements where appropriate.

Section I: Replaced conversion tables with conversion charts. AddedU.S. gallons to liters to imperial gallons conversion chart.

Section II: Table 2−1. Added R.T. and TS−1 as approved fuels.Paragraph 2−5. Updated Figure 2−5.

Section IV: Paragraph 4−12. Added requirement to turn off electricalpower before entering IIDS �TIME SUMMARY" menu.

Section V: Paragraph 5−2. Revised noise characteristics.

Paragraph 5−8 and all AEO hover charts. Removed references to 5000FT HD limit when operating at weights between 6251 LB and 6500 LBlimit.

Section VI: Table 6−1. Updated aft longitudinal CG limit at 6500 LB.Figure 6−1. Revised �Chart B".

Section VIII: Paragraph 8−11. Added battery removal requirement.Added Resetting IIDS Time/Date procedure.

Section X: Paragraph 10−7. Corrected weight limitation on landinggear. Figure 10−6. Added 6500 lb to CG Envelope.

ROTORCRAFT FLIGHT MANUAL



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Reissue 1

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CSP−902RFM207E−1ROTORCRAFT FLIGHT MANUALMD900 (902 Configuration with PW 207E) General

Original 1−i/( 1−ii blank)Reissue 1

S E C T I O N IGENERAL

TABLE OF CONTENTS

PARAGRAPH PAGE1−1. Introduction 1−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−2. Scope 1−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−3. Organization 1−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−4. Method of Presentation 1−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−5. Definition of Terms 1−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−6. Abbreviations 1−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−7. Multi−Purpose Utility Operations 1−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−8. Technical Publications 1−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−9. Rotorcraft Certification 1−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−10. Pilot’s Briefing 1−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−11. Dimensions 1−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−1. MD Explorer Rotorcraft Principal Dimensions 1−10. . . . . . . . . . . . . . . . .

Figure 1−2. Interior Dimensions and Volumes 1−11. . . . . . . . . . . . . . . . . . . . . . . . . . . .

1−12. Conversion Charts 1−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−3. Speed: MPH/Knots/KmH 1−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−4. Speed: Knots − Meters/Second 1−13. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−5. Temperature Conversion Chart 1−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−6. Liquid Measure �−� U.S. Gallons to Liters to Imperial Gallons 1−15. .

Figure 1−7. Linear Measure �−� Inches to Centimeters 1−16. . . . . . . . . . . . . . . . . . . .

Figure 1−8. Linear Measure �−� Meters to Feet 1−17. . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−9. Weight �−� Pounds to Kilograms 1−18. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 1−10. Conversion Chart: Inches of Mercury − Millibars 1−19. . . . . . . . . . . . .

Table 1−1. Standard Atmosphere Table 1−20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


MD900 (902 Configuration with PW 207E) General

Original 1−1Reissue 1

SECTION IGENERAL

1−1. INTRODUCTION

The Rotorcraft Flight Manual has been prepared to provide the pilot with all informa-tion necessary to accomplish the intended mission with the maximum amount ofefficiency and safety.

1−2. SCOPE

This manual meets all FAA requirements for APPROVED DATA and that datais so designated.

MD Helicopters, Inc. has included additional supplemental data which is intendedto provide the pilot with information that expands, enhances and eases his task.

1−3. ORGANIZATION

This manual is organized in the following manner:

FRONT MATTER:

Contains: Log of Revisions by Date, Table of Contents, Summary of Revisions,and the List of Effective Pages.

By referring to the Log of Revisions By Date, the pilot may review a chronologicallisting of changes to the Flight Manual.

Reading the Summary of Revisions will inform the pilot of what changes havebeen made by paragraph reference. This Summary contains only the latest FlightManual Change.

The List of Effective Pages allows the pilot quick reference to page numbersand their respective revision number. The pages listed should reflect the revisionnumber that appears at the bottom of each page.

SECTION I �−� GENERAL

Information of general interest to the pilot, owner or operator of the aircraftand general rotorcraft information and conversion charts.

SECTION II �−� LIMITATIONS (FAA Approved)

Specifically defines the limiting factors, procedures and parameters within whichthe rotorcraft may be operated. FAA regulations require that limitations notbe exceeded.


MD900 (902 Configuration with PW 207E)General

Original1−2Reissue 1

SECTION III �−�EMERGENCY AND MALFUNCTION PROCEDURES(FAA Approved)

Problems which could be encountered in flight are defined and the proceduresnecessary to cope with or alleviate them are discussed. The data is recommendedby the manufacturer.

SECTION IV �− NORMAL PROCEDURES (FAA Approved)

Normal operating procedures from preflight through shutdown. The data givenis that recommended by the manufacturer.

SECTION V �−� PERFORMANCE DATA (FAA Approved)

Aircraft performance as defined within certain conditions, such as airspeed,weight, altitude, temperature, humidity, and wind velocity. Data is provided intabular or graph form to allow the pilot to determine the aircraft’s capabilitiesin relation to the intended mission and prevailing conditions.

SECTION VI �−� WEIGHT AND BALANCE DATA

Provides aircraft weight and balance operational data in chart and table formand provides examples that allow the pilot to accurately determine the aircraft’sgross weight, and whether the load is within longitudinal and lateral centerof gravity limits. Also contained in this section are the original weight and balancereport and equipment list (equipment both required and optional) installed onthe aircraft at the time of licensing.

SECTION VII − SYSTEMS DESCRIPTION

Offers a pilot−oriented technical description of the operation of each systeminstalled on the helicopter.

SECTION�VIII −�AIRCRAFT HANDLING, SERVICING, MAINTENANCE AND TESTING

The information contained in this section is extracted from the Handbook ofMaintenance Instructions and is highly selective. The subjects chosen are thosewith which the pilot may have direct involvement either while at his normalbase of operations or in the field.

SECTION IX − ADDITIONAL OPERATIONS AND PERFORMANCE DATA

The information provided in Section IX is given by the manufacturer to furtherassist the pilot in obtaining maximum utilization of the rotorcraft. It also providesthe pilot with abbreviated checklists as well as additional performance data.

SECTION X OPTIONAL EQUIPMENT (FAA Approved)

Certain optional equipment is available for performance of specific tasks. Inmany cases the equipment is removable and may be used in combination(s) withother optional items. Whenever the installation of an option affects FAA approvedlimitations, normal/emergency procedures or performance (Sections II thru V),an FAA approval is required. In addition, a tabular listing of all options is providedas well as a table showing the compatibility of the various options with oneanother.




SECTION XI Category A Operations (FAA Approved)

Information contained in this section pertains to Category A operations onlyand supplements information that appears in Sections I thru X of this manual.

At the front of each section there is an table of contents that lists the data by para-graph number, title, and page number.

1−4. METHOD OF PRESENTATION

General information in the various sections is provided in narrative form. Otherinformation is given in step−by−step procedures, graphs, charts, or tabular form.

The information in the step−by−step procedure is presented in the imperative mode;each statement describing a particular operation to be accomplished. Expansionof the steps is accomplished as follows:

A black change bar ( l ) in the page margin designates the latest new or changed

information appearing on that page. A hand points to changes in the contents

of an illustration.

A WARNING brings to the pilot’s immediate attention thatequipment damage and/or personal injury will occur if theinstruction is disregarded − placed after the instruction/step.

A CAUTION alerts the individual that equipment damage may resultif the procedural step is not followed to the letter − placed afterthe instruction/step.

NOTE: A NOTE expands upon and explains the preceding step and provides fullerunderstanding of the particular operation.

1−5. DEFINITION OF TERMS

The concept of procedural word usage and intended meaning has been adheredto in preparing this manual is as follows:

�Shall" has been used only when the application of a procedure is mandatory.

�Should" has been used only when the application of a procedure is recommended.

�May" and �need not" have been used only when the application of a procedureis optional.

The terms IMMEDIATELY, POSSIBLE, and PRACTICAL as used in this manualrefer to the degree of urgency with which a landing must be made.

LAND IMMEDIATELY − Execute a power−on approach and landing without delay.

WARNING

CAUTION




LAND AS SOON AS POSSIBLE − Execute a power−on approach and landingto the nearest safe landing area that does not further jeopardize the aircraft oroccupants.

LAND AS SOON AS PRACTICAL − Extended flight is not recommended. Whetherto complete the planned flight is at the discretion of the pilot−in−command. However,the nature of the specific problem or malfunction may dictate termination of theflight before reaching the destination.

1−6. ABBREVIATIONS

SIGNS

> Greater than

� Equal to or greater than

< Less than

� Equal to or less than

A

AC Air Conditioner

AEO All Engines Operating

A/N Alphanumeric

AGL Above Ground Level

ALT Alternate; Altitude

AOG Aircraft On Ground

APU Auxiliary Power Unit

ASCM Aircraft Systems ConditionMonitoring

ATT Attitude

B

BAT Battery

BIT Built In Test

BL Butt Line

BLD Bleed

BMS Balance MonitoringSystem

C

CAB Cabin

CAB HEAT Cabin Heat

CC Cubic Centimeter

CCW Counter Clockwise

CKP(T) Cockpit

CLP Collective Lever Position

Cm Centimeters

COM Communication

CW Clockwise

D

dBA A−weighted Decibel

DIR Direction; Directional

E

ECS Environmental ControlSystem

ECTM Engine Condition TrendMonitoring

EEC Electronic Engine Control

EGT Exhaust Gas Temperature

ENG Engine

ESNTL Essential

ETL Effective Translational Lift

EXT Extend; External




F

FAA Federal AviationAdministration

FADEC Full Authority DigitalElectronic Control

FAR Federal AviationRegulation

FMU Fuel Metering Unit

FSO Flights since overhaul

Ft Feet

Ft/Min Feet per Minute

FWD Forward

G

GA Go−around

GCU Generator control unit

GEN Generator

GBMC Ground−basedMaintenance Computer

GPU Ground Power Unit

H

HAT Height Above Touchdown

HD Density Altitude

Hg Mercury

HIRF High Intensity RadiatedField

HP Pressure Altitude

HSI Horizontal SituationIndicator; Hot SectionInspection

HVR Hover

HYD Hydraulic

I

IAS Indicated Airspeed

ICS Intercom System

IFR Instrument Flight Rules

IGE In Ground Effect

IIDS Integrated Instrumentation

Display System

IMC Instrument Meteorological

Conditions

INST Instrument

IPS Inlet Particle Separator

In Inches

INST(R) Instrument

IVSI Instantaneous Vertical

Speed Indicator

K

Kg KG Kilogram

KIAS Knots Indicated Airspeed

Km KM Kilometer

Km/H KM/H Kilometers per

Hour

KT Knots

KTAS Knots True Airspeed




L

L Left; Liters

LB Pound

Lb(s) Pound(s)

L.H. Left Hand

LND Landing

LT Light

M

M m Meters

MBAR Millibar

MCP Maximum ContinuousPower

Min Minutes

MPH Miles−Per−Hour

M/R Main Rotor

MSTR Master

N

NAV Navigation

NG Gas Producer RPM

NP Power Turbine RPM

NR Rotor Speed

O

OAT Outside Air Temperature

OEI One Engine Inoperative

OGE Out of Ground Effect

OVRD Override

OVSP Overspeed

P

PLA Power Lever Angle

PMA Permanent MagnetAlternator

PNL Panel

POSN Position

PRI Primary

Pt Pint

PWC Pratt and Whitney Canada

R

R Right

REL Release

RET Retract

R.H. Right Hand

RTR Rotor

S

Sec Seconds

SEL Sound Exposure Level

SL Sea Level

SLT Searchlight

SSO Starts since overhaul

STA Station

STBY Standby

STC Supplemental TypeCertificate

SYS System

T

TBO Time Between Overhaul

TOP Takeoff Power

TSN Time Since New

TSO Time Since Overhaul




U

U.S. gal U.S. gallons

V

VFR Visual Flight Rules

VH Maximum speed in levelflight at MCP

VLV Valve

VMC Visual MetrologicalConditions

VNE Never Exceed Speed

Vs Versus

VSCS Vertical StabilizationControl System

VY Best Rate of Climb Speed

W

WL Water Line

X

XFD Crossfeed

XMSN Transmission

XPNDR Transponder

1−7. MULTI−PURPOSE UTILITY OPERATIONS

The installation and use of certain optional equipment is approved by the FAAand requires supplemental flight data when limitations, performance or proceduresare affected. Refer to Section X for Optional Equipment.

MD Helicopters, Inc. optional equipment items and STC items which are FAA ap-proved for the MD EXPLORER may be installed and used.

1−8. TECHNICAL PUBLICATIONS

A file of technical publications is available to aid in obtaining maximum utilizationof your rotorcraft. Revisions and new issue publications are provided to continuallyupdate and expand existing data.

MDHI Publications Revisions and Reissues

Changes in limitations, procedures, performance, optional equipment, etc., re-quire flight manual revisions and change or replace flight manual content asappropriate. To ensure that MDHI manuals continue to show current changes,revised information is supplied as follows.

Revisions

Change to parts of the manual by the replacement, addition and/or deletionof pages is done by revision. The List of Effective Pages that accompanieseach revision, identifies all affected pages. Such pages must be removed fromthe manual and discarded. Added or replaced pages must be put in and ex-amined against the List of Effective Pages.




Reissues

Occasionally the manual may be reissued and is identified as ‘‘Reissue #1,Reissue #2’’, etc. The preceding issue of the manual then becomes obsoleteand must be discarded. The reissue includes all prior revisions. All pagesin a reissue become ‘‘Original’’ pages. The reissue may also include new orchanged data. These changes will be identified on the ‘‘Summary of Revisions’’page as well as having change bars appear in the page margin on the effectedpages.

The following publications are available.

Rotorcraft Flight Manual (RFM).

Rotorcraft Maintenance Manual (RMM)

Servicing and Maintenance

Instruments − Electrical − Avionics

Component Maintenance Manual (CMM)

Structural Repair Manual (SRM)

Illustrated Parts Catalog (IPC)

Service Information Bulletins and Letters

New and revised publications are available through MDHS Subscription Service.Further information may be obtained by contacting:

MD Helicopters, Inc.M615−G0484555 E McDowell RdMesa, AZ 85215−9734

or your local Service Center, Distributor, or Sales Company.

All persons who fly or maintain MD helicopters are urged to keep abreast of thelatest information by using the subscription service.




1−9. ROTORCRAFT CERTIFICATION

Certified under FAR Part 27 through amendment 27−26 dated April 5, 1990, SpecialCondition for High Intensity Radiated Fields (HIRF) protection per FAR 21.16; FARPart 36 Appendix J, Noise, effective on the date of Type Certification, and FARPart 27 Appendix C Criteria for Category A effective August 8, 1996.

The rotorcraft is certified by the Federal Aviation Administration under FAA TypeCertificate Number H19NM.

The FAA model designation is MD900

The FAA/ICAO aircraft type designator is EXPL

The MD Helicopters, Inc. commercial designation is MD Explorer

1−10.PILOT’S BRIEFING

Prior to flight, passengers should be briefed on the following.

Approach and depart the rotorcraft from the front in full view of the pilot, beingaware of the main rotor.

Use of seat belts and shoulder harnesses.

Smoking.

The opening and closing of doors.

Evacuation of the aircraft in an emergency.

Location and use of emergency/survival equipment.




1−11.DIMENSIONS

Refer to Figure 1−1 and Figure 1−2 for exterior dimensions and interior volumes.

F92−002B

5.33 FT

5 0’°

9.17 FT

STATIC GROUND LINE

@ DESIGN GROSS WEIGHT

3 16’°

12.00 FT

10.92 FT

9.33 FT

7.33 FT

33.83 FT

(1.62 M)

5.92 FT(1.80 M)

(10.34 M)

(2.23 M)

(3.66 M)

(3.33 M)

(2.79 M)

40.58 FT(12.37 m)

34.08 FT(10.39 M)

Figure 1−1. MD Explorer Rotorcraft Principal Dimensions




4.08 FT (1.2 M)

4.16 FT (1.2 M) WITH DOOR ON4.33 FT (1.3 M) WITH DOOR OFF

ENTIRE AFT CABIN172.5 FT3 (4.9 M3)

BAGGAGE COMPARTMENT51.4 FT3 (1.5 M3)

4.75 FT (1.4 M)

18.25 FT (5.5 M)

12.9 FT (3.9 M)

6.25 FT (1.9 M)

F90−003

Figure 1−2. Interior Dimensions and Volumes




1−12.CONVERSION CHARTS

F92−004

200

180

160

140

120

100

80

60

40

20

0

180160140120100806040200

200

180

160

140

120

100

80

60

40

20

0

220

240

260

280

300

320

MP

H

Km

/H

KNOTS

EXAMPLE: CONVERT 100 KNOTS TO MPH AND TO KM/HR: ENTER CHART AT 100 KNOTS AND FOLLOW ARROW TO SLOPING LINE. TO FIND MPH, MOVE LEFT AND READ 115 MPH. TO FIND KM/HR, MOVE RIGHT FROM THE SLOPING LINE AND READ 185 KM/HR

Figure 1−3. Speed: MPH/Knots/KmH




F92−006

METERS/SECKNOTS

CONVERT KNOTS TO METERS/SEC

KNOWN:WIND SPEED = 25 KT

METHOD:ENTER CHART AT 25 KT READAPPROXIMATELY 13 METERS/SECACROSS ON METERS/SEC SCALE

METHOD MAY BE REVERSED TO FIND KNOTS WHEN METERS/SEC AREKNOWN

EXAMPLE

0

5

10

15

20

25

30

35

40

45

50

0

5

10

15

20

25

Figure 1−4. Speed: Knots − Meters/Second




CONVERT °F TO °C

KNOWN:TEMPERATURE = 50° F

METHOD:ENTER AT 50° FREAD 10° C ACROSS ON °C SCALE

METHOD MAY BE REVERSED TO FIND ° F WHEN ° C IS KNOWN

ALTERNATE METHOD:° F = (9/5 X °C) + 32°C = 5/9(°F − 32)

TEMPERATURE

EXAMPLE:

°°F C140

120

100

80

60

40

20

0

−20

−40

−60

60

40

20

0

−20

−40

−60−80

F92−005

−50

−30

−10

10

30

50

Figure 1−5. Temperature Conversion Chart




F927−150

U.S. GALLONS

1901801701601501401301201101009080706050403020100

LIT

ER

S

IMPERIAL (BRITISH) GALLONS

700

650

600

550

500

450

400

350

300

250

200

150

100

50

0

0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150

Figure 1−6. Liquid Measure − U.S. Gallons to Liters to Imperial Gallons




F927−152

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

29

30

31

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0 9.5 10.0 10.5 11.0 11.5 12.0

INCHES

CE

NT

IME

TE

RS

Figure 1−7. Linear Measure − Inches to Centimeters




F927−151

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 65000

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

FE

ET

METERS

20000

Figure 1−8. Linear Measure − Meters to Feet




F927−149

1900

1800

1700

1600

1500

1400

1300

1200

1100

1000

900

800

700

600

500

400

300

200

100

0

0 500 1000 1500 2000 2500 3000 3500 4000

3000

2900

2800

2700

2600

2500

2400

2300

2200

2100

2000

1900

1800

POUNDS

KIL

OG

RA

MS

4500 5000 5500 6000 65004000

Figure 1−9. Weight − Pounds to Kilograms




F92−007

MILLIBARS

1000995990985980975970965960955950945

29.5

29.4

29.3

29.2

29.1

29.0

28.9

28.7

28.6

28.5

28.4

28.3

28.2

28.1

28.0

28.8

10551050104510401035

EXAMPLE 1: 29.44 IN. Hg = 997 mbarEXAMPLE 2: 30.18 IN. Hg = 1022 mbar

10351030102510201015101010051000

IN. H

g

30.6

30.7

30.8

30.9

31.0

31.1

30.5

2

1 30.5

29.5

29.6

29.7

29.8

29.9

30.0

30.1

30.2

30.3

30.4

Figure 1−10. Conversion Chart: Inches of Mercury − Millibars




Table 1−1. Standard Atmosphere Table

Standard Sea Level Conditions:Temperature: 59°F (15°C)Pressure: 29.921 in.Hg (1013.25 mbar)Density: 0.0023769 slugs/ft3 (1.225 kg/m3)

ALTITUDE(feet)

DENSITYRATIO σ

1 �σ

TEMPERATURE PRESSURE(mbar)

PRESSURE(in. Hg)

PRESSURERATIO(°C) (°F)

0 1.0000 1.000 15.00 59.000 1013.25 29.921 1.0000

1000 0.9711 1.0148 13.019 55.434 997.18 28.856 0.9644

2000 0.9428 1.0299 11.038 51.868 942.14 27.821 0.9298

3000 0.9151 1.0454 9.056 48.302 908.14 26.817 0.8962

4000 0.8881 1.0611 7.076 44.735 875.12 25.842 0.8637

5000 0.8617 1.0773 5.094 41.196 843.08 24.896 0.8320

6000 0.8359 1.0938 3.113 37.603 811.99 23.978 0.8014

7000 0.8106 1.1107 1.132 34.037 781.86 23.088 0.7716

8000 0.7860 1.1279 −0.850 30.471 752.63 22.225 0.7428

9000 0.7620 1.1456 −2.831 26.905 724.29 21.388 0.7148

10000 0.7385 1.1637 −4.812 23.338 696.82 20.577 0.6877

11000 0.7155 1.1822 −6.793 19.772 670.21 19.791 0.6614

12000 0.6932 1.2011 −8.774 16.206 644.40 19.029 0.6360

13000 0.6713 1.2205 −10.756 12.640 619.44 18.292 0.6113

14000 0.6500 1.2403 −12.737 9.074 595.23 17.577 0.5875

15000 0.6292 1.2606 −14.718 5.508 571.83 16.886 0.5643

16000 0.6090 1.2815 −16.669 1.941 549.14 16.216 0.5420

17000 0.5892 1.3028 −18.680 −1.625 527.23 15.569 0.5203

18000 0.5669 1.3246 −20.662 −5.191 505.99 14.942 0.4994

19000 0.5511 1.3470 −22.643 −8.757 485.48 14.336 0.4791

20000 0.5328 1.3700 −24.624 −12.323 465.63 13.750 0.4595

21000 0.5150 1.3935 −26.605 −15.899 446.47 13.184 0.4406

22000 0.4976 1.4176 −28.587 −19.456 427.91 12.636 0.4223

23000 0.4806 1.4424 −30.568 −23.002 409.99 12.107 0.4046

24000 0.4642 1.4678 −32.549 −26.588 392.72 11.597 0.3874

25000 0.4481 1.4938 −34.530 −30.154 375.99 11.103 0.3711


FAA ApprovedReissue 1Original 2−i/(2−ii blank)

S E C T I O N I ILIMITATIONS

TABLE OF CONTENTS

PARAGRAPH PAGE2−1. Flight Restrictions 2−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−2. Environmental Operating Conditions 2−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−1. Ambient Temperature Envelope 2−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−2. WAT Limit and �Area A" Azimuth For Crosswind Operations 2−3. . .

2−3. Airspeed Limitations 2−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−3. VNE Chart 2−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−4. Weight Limitations 2−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−4. Minimum Flying Weight 2−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−5. Center of Gravity (CG Envelope) 2−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−5. Center of Gravity Envelope 2−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−6. Rotor Brake Limitations 2−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−7. Rotor Speed Limitations 2−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−8. Transmission Limitations 2−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−9. Power Plant Limitations 2−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−10. Generator Limitations 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−11. Starter limitations 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−12. Fuel System Limitations 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 2−1. Fuel Specifications 2−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Figure 2−6. Primary IIDS Display 2−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−7. NP and NR Scales 2−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−8. Engine Torque 2−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−9. Engine Exhaust Gas Temperature 2−11. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−10. Secondary IIDS Display 2−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−11. Engine Display 2−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−12. Transmission and Fuel Quantity Display 2−13. . . . . . . . . . . . . . . . . . . .

Figure 2−13. Airspeed Indicator 2−13. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2−14. Decals and Placards 2−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−14. Decals and Placards (Sheet 1 of 2) 2−14. . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 2−14. Decals and Placards (Sheet 2 of 2) 2−15. . . . . . . . . . . . . . . . . . . . . . . . . . .


Limitations



FAA ApprovedReissue 1Original

2−1

SECTION IILIMITATIONS

2−1. FLIGHT RESTRICTIONS

Approved as an eight place (maximum) helicopter.The minimum flight crew consists of one pilot operating the helicopter from the rightseat. The left crew seat may be used for an additional pilot when the approved dualcontrols are installed.Under seat baggage stowage:

Placing of cargo or baggage under seats (including crew seats) is permitted onlywhen the seat is unoccupied.

Aerobatic flight:Aerobatic flight is not allowed.

Aircraft equipped with Bendix/King KFC900 Flight Control System:NOTE: The following information supersedes applicable limitations found in Bendix/King IFR

Avionics/KFC 900 RFMS 006−00845−0000 and 006−00845−0004 for STCSR00436WI−D.

For VFR flights at gross weights between 6251 and 6500 lb (2835 and 2948kg):

Maximum airspeed with autopilot engaged is 100 KIASMaximum Operating Altitude with autopilot engaged 5000 FT HD

For IFR flights at gross weights between 6251 and 6500 lb (2835 and 2948 kg):Autopilot must be operational.Maximum airspeed with autopilot engaged is 100 KIASMaximum Operating Altitude with autopilot engaged 5000 FT HD

Flight with doors opened or removed is approved under the followingconditions.

Baggage door removed:With the baggage door removed and cockpit and cabin doors closed, maximumairspeed is limited to the 140 KIAS envelope shown in Figure 2−3.

Approved doors off configurations:Maximum airspeed is limited to the 100 KIAS envelope shown in Figure 2−3.Both cockpit doors removedBoth cabin doors removedBoth cockpit and both cabin doors removed

Cabin doors open in flight:Maximum airspeed is limited to 60 KIAS (with or without cockpit doors).One or both cabin doors may be opened or closed in flight at airspeeds up to60 KIAS.Maximum airspeed is limited to 100 KIAS (with or without cockpit doors) follow-ing installation of modified upper door fittings (Ref. Figure 2−14).For sustained flight with the cabin doors open, use of the cabin door hold opendevice is required.

NOTE: Baggage compartment door may be removed with any of the above configurations.


MD900 (902 Configuration with PW 207E)Limitations


2−2

2−2. ENVIRONMENTAL OPERATING CONDITIONS

Kinds of Operations:

This rotorcraft is certified in the normal helicopter category for day and nightVFR operation when the appropriate instruments and equipment required bythe airworthiness and/or operating rules are approved, installed and are in oper-able condition.

Maximum operating altitude at gross weights 6250 lb (2835 kg) and be-low: 20,000 Feet HD

Maximum operating altitude at gross weights 6251 to 6500 lb (2835 to2948 kg): 14,000 Feet HD.

Maximum altitude for HIGE/takeoff and landing operations: Refer toFigure 2−2.

F927−001C

20000

18000

16000

14000

12000

10000

8000

6000

4000

2000

0

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

FREE AIR TEMPERATURE − °C

PRIMARY FUELSONLY

PRIMARY ANDSECONDARY

FUELS

DASHED LINES FORAIRCRAFT WITHOUTGENERATOR COOLINGMODIFICATION

−20 −10 0 10 20 30 4052

50−30−40−50−36

HIGE OPERATIONSLIMITED TO 5 MINUTESAT TEMPERATURESFROM 50 TO 52°C

40.6

14000HD LIMIT FORGROSS WEIGHTSFROM 6251 TO 6500 LB

Figure 2−1. Ambient Temperature EnvelopeIIDS Built In Test − cold temperature:

A commanded IIDS BIT must be performed prior to the first start of the dayif the helicopter has been statically exposed to temperatures below 0°C for 12hours or longer.

NOTE: The IIDS display may not be readable during the initial power up BIT whenstatically exposed to the above ambient temperatures.

Icing conditions:

Flight into known icing conditions is prohibited.


Limitations




2−3

Snow conditions (IPS installed):

Flight into falling or blowing snow is only permitted when the NACA inlet switchis in the closed position. The switch shall remain in the closed position for theduration of the flight, even after leaving the falling or blowing snow conditions.

Cabin heat:Cabin heat must be OFF in the crew and passenger compartments when ambienttemperatures are greater than 28°C (82°F).

F927−002C

ÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏÏ

GROSS WEIGHT − LBS

DE

NS

ITY

ALT

ITU

DE

− F

EE

T

120°

135°

AZIMUTH RANGE FOR AREA A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

4000 4200 4400 52005000 60006500

4600 4800 5400 5600 5800 6200 6400 66006250

12400

OBSERVE THE MAXIMUM WEIGHT,

ALTITUDE, TEMPERATURE (WAT)

LIMITS FOR TAKEOFF AND LANDING

IGE HOVER OPERATIONS HAVE BEENDEMONSTRATED IN WINDS UP TO 17 KNOTSFROM ANY AZIMUTH.

IGE HOVER OPERATIONS HAVE BEENDEMONSTRATED IN WINDS UP TO 17 KNOTS FROMALL AZIMUTHS EXCEPT BETWEEN 120� AND 135� ANDIGE HOVER OPERATIONS HAVE BEENDEMONSTRATED IN WINDS UP TO 15 KNOTS FORAZIMUTHS BETWEEN 120� AND 135�.

AVOID STEADY IGE HOVER OPERATIONS IN AREA A WHEN WINDS AREGREATER THAN 15 KNOTS FROM AZIMUTHS BETWEEN 120� TO 135�.

MAXIMUM SAFE WINDS FOR HOVER OPERATIONS DECREASE WITHINCREASING DENSITY ALTITUDE. TAKEOFF AND LANDING OPERATIONS INCALM WINDS OR HEADWINDS

Figure 2−2. WAT Limit and “Area A” Azimuth For Crosswind Operations




2−4

2−3. AIRSPEED LIMITATIONS

Observe gross weight depictions on chart.

VNE is 60 KIAS with lateral C.G. greater than +2 inches.

VNE is 134 KIAS at 6500 lb (2948 kg) following compliance with SB900−105.

VNE decreases at a rate of 4 kts/1000 FT above 5500 Feet HD

ÏÏÏÏÏÏÏÏÏÏÏÏ

20000

15000

10000

5000

0

40 50 60 70 80 90 100 110 120 150130 140

DE

NS

ITY

ALT

ITU

DE

− F

EE

T

INDICATED AIRSPEED − KNOTS

−30°C

−36°C

F92−010B

VNE: AUTOROTATIONOEI OPERATIONS

HYDRAULICS FAILUREVSCS FAILURE

XMSN CHIP LIGHT

140 KIAS ENVELOPE

<5100 LB

VNE : POWER ONVNE: � −25°C

5100 − 6250LB

: VNE DUE TOTIP MACH

: HD AND VNELIMIT FOR 6251TO 6500LB

100 KIASENVELOPE

134

Figure 2−3. VNE Chart

2−4. WEIGHT LIMITATIONS

Serial numbers 900−00117 and subsequent: Maximum gross weight 6500 lb (2948 kg).

Serial numbers 900−00116 and prior:

If SB900−099R1 and SB900−102R1 have been accomplished: Maximum grossweight 6500 lb (2948 kg).

If only SB900−099R1 has been accomplished: Maximum gross weight 6250 lb(2835 kg).

If SB900−099R1 and SB900−102R1 have not been accomplished: Maximum grossweight 5400 lb (2449 kg).

Minimum flying gross weight: Refer to Figure 2−4.


Limitations




2−5

Cargo deck capacity: 1500 lb. (680 kg) not to exceed 115 lb/ft2 (4.85 kg/m2).

Maximum weight in baggage compartment (sta. 234 to 257): 500 lb. (227 kg) notto exceed 115 lb/ft2 (21.07 kg/m2).

F92−169A

DE

NS

ITY

ALT

ITU

DE

− F

EE

T

GROSS WEIGHT − LB

20000

4000

2000

0

−2000

−4000

−6000

−8000

3000 3500 4000

1153

−6812

4185

Figure 2−4. Minimum Flying Weight

2−5. CENTER OF GRAVITY (CG ENVELOPE)

Ensure helicopter CG and weight are within approved limits throughout flight.

Expanded lateral C.G.:

Maximum lateral C.G. for takeoffs and landings from/to a surface is + 2 inches.

VNE is 60 KIAS with lateral C.G. greater than +2 inches.

Longitudinal C.G. envelope is as shown on chart �B" below when lateral C.G.is greater than +2 inches.

F92−011B

3000

3500

4000

4500

5000

5500

6000

6500

−3 −2 −1 0 1 2 3 4 5 6 73000

3500

4000

4500

5000

5500

6000

6500

194 196 198 200 202 204 206 208

GR

OS

S W

IGH

T − L

BSG

RO

SS

WIG

HT

− L

BS

WHEN OPERATING IN THEEXPANDED CG REGION OFCHART A, THE MAXIMUMLONGITUDINAL C.G. LIMIT,AS DEPICTED BY THEDASHED LINE IN CHART B,APPLIES.

CHART A: LATERAL C.G. STATION (IN.)

EXPANDEDCG LIMITS

CHART B: LONGITUDINAL C.G. STATION (IN)

5100 LBS

Figure 2−5. Center of Gravity Envelope




2−6

2−6. ROTOR BRAKE LIMITATIONS

The rotor brake must be in the stowed position prior to engine starting.

The rotor brake may be applied after both engines are shutdown with NR at orbelow 70 percent.

2−7. ROTOR SPEED LIMITATIONS

Power on:

Continuous operation Maximum 101%Minimum 99%

Transient Range: 91% to 98%102% to 108%

Power off:

Continuous operation: 108% maximum88% minimum

2−8. TRANSMISSION LIMITATIONS

Maximum transmission oil pressure: 104% PSI

Minimum transmission oil pressure: See Figure 2−12.

Maximum transmission oil temperature: 110°C

Minimum transmission oil temperature: −18°C


Limitations




2−7

2−9. POWER PLANT LIMITATIONS

The pilot shall monitor the IIDS during all phases of operation andrecord and report any exceedances to maintenance as soon aspossible.

Torque limits:

Normal:

Maximum continuous: 100%

Takeoff (5 minute): 101% to 110%

Maximum transient over torque: 111% to 124% for 10 seconds

OEI limits:

Maximum continuous: 124%

2.5 minute: 125% to 135%

Torque greater than 135%: NOT ALLOWED

Exhaust gas temperature limits:

Normal limits:

Maximum continuous: 850°CTakeoff (5 minutes): 851°C to 900°CTransient limits: 901°C to 1000°C for 20 seconds

OEI limits:

Maximum continuous: 900°C2.5 minute: 901°C to 970°CMaximum transient limits: 971°C to 1000°C for 20 seconds

Overtemperature limits for starting:

Maximum 875°C for 2 seconds711°C for 10 seconds650°C for 45 seconds

NOTE: Engine start should be completed within 45 seconds with EGT stabilized below650°C.

CAUTION




2−8

Output shaft (NP) speed limits:

Normal operating range: 99% to 101%

Transient limits: >102% to 108% for 20 seconds (not cumulative)

NG limitations:

Normal limits:

Maximum continuous: 97.2%

Takeoff (5 minutes): 97.3% to 99.8%

Transient limits: 99.9% to 104.1% for 20 seconds

OEI operating limits:

Maximum continuous: 99.8%

2.5 minute: 99.9% to 103.0%:

Transient limits: 103.1% to 104.1% for 20 seconds

Engine oil system limitations:

Engine oil temperature limits:

During starting: −40°C to 125°CNormal operating range: 10°C to 120°C

Engine oil pressure limits:

Normal operating range: 85% to 100% psi

Maximum: 100% psi (>5 minutes)

Minimum: <80% psi (>5 seconds)


Limitations




2−9

2−10.GENERATOR LIMITATIONS

Maximum continuous: 99% for each generator.

2−11.STARTER LIMITATIONS

30 seconds on, 30 seconds off; 30 seconds on, 30 seconds off; 30 seconds on, 30 minutesoff.

2−12.FUEL SYSTEM LIMITATIONS

Table 2−1. Fuel Specifications

PRIMARY FUELS (1) SECONDARY FUELS (2)

Jet A (ASTM D1655) Jet B (ASTM D6615)

Jet A−1 (ASTM D1655) JP−4 (MIL−DTL−5624)

JP−5 (MIL−DTL−5624) TS−1 (CIS GOST 10227) (3)(4)

JP−8 (MIL−DTL−83133)

RT (CIS Standards OrganizationGOST 10227)(4)

Notes:(1). Using these primary fuels, the engine shall operate satisfactorily throughout the

altitude/temperature envelope (Ref. Figure 2−1).

(2). Using these secondary fuels, the engine shall operate satisfactorily up to 10,000 FT(Ref. Figure 2−1).

(3). Use of TS−1 is considered to be satisfactory for occasional use only: not morethan100 hours (intermittently or continuously). If fuel is used for more than 100hours, refer to PWC engine maintenance manual (3038324).

(4). Must contain one of the following anti−ice additives at a concentration up to 0.3%by volume: Ethylene Glycol Monomethyl Ether (Ethylcellosolve, Liquid I) as definedin GOST 8313, Liquid I−M (mixture 50% Liquid I with 50% methyl alcohol) as de-fined in TU−6−10−1458, Tetrahydrofurfuryl alcohol (TGF) as defined in GOST17477 or Liquid TGF−M (mixture 50% TGF with 50% methyl alcohol ) as defined inTU 6−10−1457.

Additional fuel specifications may be found in Section VIII.

Maximum 140 KIAS with either left or right low fuel warning tickmarks ON.

During operations in temperatures of 13°F (−10°C) or colder, fuel added to the tankmust contain either anti−icing additive PFA-55MB or anti−icing additive perMIL-I-27868 or MIL-I-85470 with a minimum concentration of 0.06% by volumeand a maximum concentration of 0.15% by volume. Follow manufacturer’s instruc-tions.

WARNING




2−10

2−13.INTEGRATED INSTRUMENTATION DISPLAY SYSTEM (IIDS)

..

NPNR

EECMANFAIL

NP

EECMANFAIL

ENGOUT

TORQUE

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

ENGINE TORQUE EXHAUST GASTEMPERATURE DISPLAY

POWER TURBINESPEED DISPLAY

ROTOR SPEED DISPLAY

DISPLAY

F92−012

DIGITALDISPLAYS EGT

Figure 2−6. Primary IIDS Display

NPNR

NP

NR

99 − 101%>101%

<99%

>108%

>112%

<91%

98 − 102%

>102%

>108%

<98%

<88%

>111%

<80%F927−003

NOTE: > = GREATER THAN< = LESS THAN

NP

Figure 2−7. NP and NR Scales


Limitations




2−11

>98%

>100%

>109%

F927−004


EECMANFAIL

TORQUE OEI OPERATIONS

>135%

>124%

>122%

EECMANFAIL

TORQUENORMAL

OPERATIONS

Figure 2−8. Engine Torque

..

ENGOUT

EGT

..

ENGOUT

EGT

OEI OPERATION

F927−005

>900°C

>850°C

>845°C

>970°C

>900°C

>894°C


ONLY

NORMALOPERATIONS

Figure 2−9. Engine Exhaust Gas Temperature




2−12

SECONDARY IIDS DISPLAY

CAB HEAT

BAT HOT

BAT WRM

ROTORBRAKECABINDOOR

BAGGAGEDOOR

1 HYD 2

IIDS

FIRECHIPS

%LOAD

FUEL

GEN

LB

OAT °C

%PSI°C

FIRECHIPS

FIRECHIPS

NG

%PSI°C %PSI°C

%LOAD

NG

GEN

LEFT ENGINE PARAMETERDISPLAY TRANSMISSION PARAMETER

DISPLAYRIGHT ENGINE PARAMETER

DISPLAY

FUEL QUANTITY

ENGINE OIL

ENGINE OILPRESSURE DISPLAY

GENERATOR%LOAD DISPLAY

GAS PRODUCERTURBINE SPEED DISPLAY

TEMPERATURE DISPLAY

PRESSURE DISPLAYTRANSMISSION OIL

TRANSMISSION OILTEMPERATURE DISPLAY

F92−016

Figure 2−10. Secondary IIDS Display

FIRE

CHIPS

%LOAD

GEN

%PSI° C

NG

HIGH WARNING: >125°CHIGH CAUTION: >120°C

<80% PSI >2 SEC<80% PSI >5 SEC

CAUTION: 100% LOAD

HIGH WARNING: >103.0% NGHIGH CAUTION: >99.8% NG

LOW WARNING: <50% NG

F927−052

LOW CAUTION: <10°C WITH NG >50%<−45°C WITH NG < 50%

>100% PSI >5 MINUTES

TIME

TIME

NOTE: ‘‘>’’ = GREATER THAN‘‘<’’ = LESS THAN‘‘<’’ = EQUAL TO OR LESS THAN

Figure 2−11. Engine Display


Limitations




2−13

FUEL

LB

FIRECHIPS

%PSI°C

HIGH WARNING: >110°CHIGH CAUTION: >93°C

LOW CAUTION: <−18°C

HIGH WARNING: >104% PSIHIGH CAUTION: >100% PSI

LOW CAUTION: <75% PSI NONELOW WARNING: <65% PSI <50% PSI

LOW CAUTION: 300 LBSLOW FUEL WARNING

SEGMENT: 150 LBS

F92−018A

FLIGHT

LEFT/RIGHT LOW FUEL WARNING TICK MARKS: 97 TO 127 LBS

IDLE

NOTE: ‘‘>’’ = GREATER THAN‘‘<’’ = LESS THAN

Figure 2−12. Transmission and Fuel Quantity Display

MPH

200

150

100 80

60

40

40

60120

140

160

180200

KNOTS

AIRSPEED

100

AIRSPEED INDICATOR MARKINGS:

150 KNOTS

100 KNOTS

F92−019A

0−30kt INDICATOR UNRELIABLE

40

60120

140

160

180200

KNOTS

AIRSPEED

100

100 80 100 80

Figure 2−13. Airspeed Indicator




2−14

2−14.DECALS AND PLACARDS

SEAT ATTACHONLY

NO TIEDOWN

WARNING

THIS PANEL MUST BE SECUREDPRIOR TO JACKING, TOWING

OR FLYING THE AIRCRAFT

FUEL CELL ACCESS DOOR

WARNINGREPLACE DOOR BEFORE FLIGHT

AND JACKING AIRCRAFT

F92−020−1

LOCATED AT UPPER CABINSEAT ATTACH POINTS

LOCATED ON CABIN FLOOR

LOCATED ON BAGGAGECOMPARTMENT FLOOR

NO ARTICLES TO BESTOWED UNDER SEATS

LOCATED ON COCKPIT DOORLOWER WINDOW FRAME

1. LOCATED ON UPPER COCKPIT DOOR FRAME

2. LOCATED ADJACENT

TO COCKPIT DOORACCESS HANDLE

PRIOR TO FLIGHT1. TURN HANDLE TO SAFELOCK POSITION

2. FASTEN SEAT BELTS AND SHOULDER HARNESS

SLIDING

DOOR

EMERGENCY EXIT

PULL TAB

TO REMOVE WINDOW

NO ARTICLES TO BESTOWED UNDER SEATS

LOCATED ON CABIN DOORUPPER FRAME (FWD)

LOCATED ON CABIN DOORUPPER FRAME (CENTER)

LOCATED ON CABIN DOORUPPER FRAME (AFT)

LOCATED ON CABIN DOOR FRAMEADJACENT TO EMERGENCY EXITRELEASE HANDLE

Figure 2−14. Decals and Placards (Sheet 1 of 2)


Limitations




2−15/(2−16 blank)

DURING OPERATIONS IN TEMPERATURES OF 13�F (−10�C)OR COLDER, FUEL ADDED TO THE TANK MUST CONTAIN

EITHER ANTI-ICING ADDITIVE PFA-55MB OR ANTI-ICINGADDITIVE PER MIL-I-27868 OR MIL-I-85470 WITH AMINIMUM CONCENTRATION OF .06% BY VOLUME

AND A MAXIMUM CONCENTRATION OF .15% BY VOLUME.SEE FLIGHT MANUAL FOR MIXING PROCEDURES.

F92−020−2B

STATIC PORTKEEP HOLES ANDSURFACE CLEAN

NO STEP

USEABLE CAP. 158.5 U.S. GALS.

LOCATED ABOVE FUEL FILLER

LOCATED ABOVE FUEL FILLER

LOCATED ON FILLER NECK

LOCATED ON FILLER NECK

LOCATED ABOVE STATIC PORT

RADIO CALLN X X X X X

ROTORBRAKE

LIFT HANDLE,

ROTATE CW,

PULL DOWN

DO NOTENGAGE ROTORBRAKE ABOVE

70% NR

THIS HELICOPTER MUST BE OPERATEDIN COMPLIANCE WITH THE OPERATING

LIMITATIONS SPECIFIED IN THE FAAAPPROVED ROTORCRAFT FLIGHT MANUAL

LOCATED ON INSTRUMENT PANEL



LOCATED ADJACENT TO ROTOR BRAKE

VNE CHART: LOCATED ON INSTRUMENT PANEL

APPROVED FORVFR DAY/NIGHT

100 KTCAPABILITY

MODIFIED FITTING WITH DECAL

Figure 2−14. Decals and Placards (Sheet 2 of 2)

Emergency andMalfunction Procedures


FAA ApprovedReissue 1Original 3−i

S E C T I O N IIIEMERGENCY AND MALFUNCTION

PROCEDURES

TABLE OF CONTENTS

PARAGRAPH PAGE3−1. General 3−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−2. Caution and Warning Annunciators and Audio Tones 3−2. . . . . . . . . . . . . . . . . . . . . . .

3−3. Engine Emergencies 3−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Single Engine Failure 3−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Second Engine Failure 3−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Rotor RPM Warning 3−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−4. Emergency Landing Procedures 3−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Water Landing − Dual Engine Failure 3−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Water Landing − OEI/AEO 3−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−5. EEC Malfunctions 3−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−1. EEC Caution/Warning Annunciators 3−8. . . . . . . . . . . . . . . . . . . . . . . . .

EEC Critical Fault 3−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EEC NonCritical Fault 3−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EEC Manual Control 3−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−6. Engine Starting − Manual 3−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−7. Engine/Aircraft Shutdown − Manual 3−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−8. Fire Emergencies 3−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Cabin Fire/Smoke 3−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Electrical Fire 3−16. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 3−2. Engine/Transmission Deck Fire Annunciators 3−17. . . . . . . . . . . . . . . . .

Engine Fire − On Ground 3−17. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Engine FIRE − During Flight 3−17. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transmission Area Fire 3−18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−9. Flight Control Malfunctions 3−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Anti−Torque Failure − General 3−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Anti−Torque Failure − Complete Loss of Thrust 3−19. . . . . . . . . . . . . . . . . . . . . . . . . . . Anti−Torque Failure − Fixed Thruster Setting 3−20. . . . . . . . . . . . . . . . . . . . . . . . . . . .


CSP−902RFM207E−1 ROTORCRAFT FLIGHT MANUALMD900 (902 Configuration with PW 207E)


3−ii

PARAGRAPH PAGEVSCS Failure 3−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−3. VSCS Indicator 3−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Hydraulic System Failure 3−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Cyclic Trim Failure 3−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Collective Friction Failure 3−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−10. Pitot/Static System Malfunction: Single or Dual Pitot Tube Installation 3−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Static System Malfunction 3−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−4. Alternate Static Source Toggle Valve 3−23. . . . . . . . . . . . . . . . . . . . . . . . .

Pitot Heat Failure 3−24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−11. Engine and Generator Malfunction Indications 3−24. . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−5. Engine and Generator Malfunction Annunciators 3−24. . . . . . . . . . . . . .

Engine High Oil Temperature 3−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Low Oil Temperature 3−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Engine High Oil Pressure 3−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Engine Low Oil Pressure 3−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Chips 3−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

NG High 3−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

NG Low 3−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Generator High Load 3−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Generator 3−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−12. Transmission Malfunction Indications 3−28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−6. Transmission Malfunction Annunciators 3−28. . . . . . . . . . . . . . . . . . . . . .

Transmission Oil Temperature High 3−28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transmission Oil Temperature Low 3−28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Transmission Oil Pressure Low 3−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Transmission Oil Pressure High 3−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transmission Chips 3−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Transmission Input Torque Split 3−30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−13. Fuel System Display Advisories 3−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−7. Fuel System Advisory Indicators 3−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Fuel Low 3−32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Fuel Boost Pump Failure 3−32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Single Fuel Probe Failure 3−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Dual Fuel Probe Failure 3−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Impending Fuel Filter bypass 3−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Shutoff Valve Malfunction 3−34. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−14. Caution and Warning Advisories 3−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−8. Caution/Warning Cluster 3−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Cabin Heat 3−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .



FAA ApprovedReissue 1Original 3−iii/(3−iv blank)

PARAGRAPH PAGEBattery Hot 3−36. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Battery Warm 3−37. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rotor Brake 3−38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Cabin Door 3−38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Baggage Door 3−38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

IIDS 3−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−15. Other Malfunction/Advisories 3−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

IIDS Failure 3−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Discharge 3−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Extinguisher Pressure Low 3−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

IPS Bypass 3−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

NACA Inlet Malfunction 3−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Rotor Speed Display Malfunction 3−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Ground power Unit door open 3−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−16. Vibrations 3−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−17. Emergency Egress 3−42. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−9. Cabin Door Emergency Exit 3−42. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3−18. Emergency Equipment 3−43. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 3−10. Emergency Fire Extinguisher 3−43. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .





FAA ApprovedReissue 1Original 3−1

SECTION III EMERGENCY AND

MALFUNCTION PROCEDURES

3−1. GENERAL

The procedures contained in this section are to be followed in the event of an emergen-cy or malfunction that may potentially affect the safety of the aircrew, passengers,aircraft, or personnel on the ground.

These procedures are recommended to minimize danger to the helicopter. However,these procedures should not limit the pilot from taking additional actions if thesituation warrants.

In the event of an emergency or malfunction, the pilot’s primary considerationis control of the aircraft. Then, the pilot must identify the problem and performthe appropriate procedures relevant to the situation.

Terms such as ‘‘land immediately’’, ‘‘land as soon as possible’’, and ‘‘land assoon as practical’’ are defined in Section I.


MD900 (902 Configuration with PW 207E)Emergency andMalfunction Procedures

FAA ApprovedReissue 1Original3−2

3−2. CAUTION AND WARNING ANNUNCIATORS AND AUDIO TONES

A red warning or yellow caution annunciator will illuminate on the IIDS displayand in some cases, an audio warning will sound announcing a failure or malfunction.

Some secondary IIDS displays have a digital display with a corresponding caution/warning annunciator. Pilots should insure that both the digital display and its ap-propriate caution/warning annunciator are in agreement before executing the properemergency procedure. If they do not agree, other parameters should be cross-checkedin an attempt to validate a given abnormal indication.

The following logic applies to the warning advisories:

1. ENG OUT, FIRE, CAB HEAT, and LOW FUEL warning tick marks flash(only go OFF if condition that caused the warning goes away).

2. All other warnings turn ON continuously (only go OFF if condition that causedthe warning goes away).

Audio warnings alert the pilot through the headset that a malfunction has occurredthat may require immediate corrective action.

The warning tone: The warning tone will sequence a high and low tone twiceto indicate a warning condition. These tones are activated for FIRE, CAB HEAT,BAT HOT, and EEC FAIL.

Low Rotor Audio tone: The low rotor RPM tone (a sweep tone) is activated forrotor RPM less than 95% and either engine operating, or activated for rotorRPM less than 88% and both engines failed. When the IIDS senses an enginefailure, the low rotor RPM tone is activated for one cycle. The audio tones aredisabled when the aircraft is on the ground.






3−3. ENGINE EMERGENCIES

ENGOUT

TORQUE EGT

The ENG OUT warning annunciator is located between the TORQUE and EGTvertical displays. When the IIDS senses an engine failure, the ENG OUT

warning flashes and the low rotor RPM tone is activatedfor one cycle. The EGT and TORQUE displays also re−scale. The ENG OUT advisory is disabled with aircrafton the ground.

SINGLE ENGINE FAILURE

Indications: ENG OUT annunciator illuminated and low rotor tone on for one cycle.Affected engine torque, NP and NG decreasing to zero.

Conditions: At a hover − IGE:

Procedures:

� Land

Conditions: At a hover − OGE:

Procedures:

� Collective pitch ADJUST TO MAINTAIN OEI LIMITS

NOTE: The decision to land or fly−away, following a single engine failure, will dependon ambient conditions and aircraft gross weight. Refer to Section V for best rateof climb speed, single engine rate of climb and descent, and height velocityenvelope performance data.




Conditions: In flight:

� Maintain operating engine within OEI limits

� Reduce airspeed to 100 KIAS or less (Ref. Section II)

Identify affected (failed) engine by cross checking torque, NP andNG prior to performing the following steps.

� Engine control switch OFF ON AFFECTED ENGINE� Fuel boost pump OFF ON AFFECTED ENGINE� Fuel shutoff valve OFF ON AFFECTED ENGINE

NOTE: With fuel shutoff valve OFF, fire extinguishing system is now armed.

NOTE: Fuel warning tick mark on side of affected engine may illuminate prior to tick markfor operating engine.

NOTE: If attempting a restart, leave fuel shutoff valve ON. Do not attempt restart if amalfunction is suspected.

� Land as soon as practical

NOTE: If cabin heat or air-conditioning is being used when an engine failure occurs itwill automatically be switched OFF to enable the pilot to utilize the maximumavailable power from the remaining engine for a safe recovery. If, afterrecovering to a safe OEI flight condition, cabin heat is needed for windscreendefogging, cabin comfort, etc., select the CAB HEAT OVRD position to restorecabin heat. Prior to the OEI landing, insure CAB HEAT is OFF to ensure that themaximum power is available from the remaining engine.

� Air start:

� Failed engine control switch TO OFF THEN TO FLY� IIDS MONITOR

CAUTION






SECOND ENGINE FAILURE

Indications: Low rotor RPM with low rotor tone ON if RPM is below 88%.Possible left yawTorque, NP, and NG decreasing to zero.

Procedures:

� Lower collective and maintain rotor speed within limits.

� Perform autorotation to suitable landing area.

LOW ROTOR RPM WARNING

Procedures:

� Adjust collective as necessary to control rotor RPM.

Respond immediately to the ENGINE OUT/low rotor RPM warningby adjusting collective to maintain rotor RPM within limits, thencheck engine instruments and other indications to confirm enginetrouble.

WARNING




3−4. EMERGENCY LANDING PROCEDURES

WATER LANDING − DUAL ENGINE FAILURE

Procedures:

� Adjust collective pitch as necessary to establish autorotation.

� Cabin doors may be opened when airspeed is 60 KIAS or less.

� Make autorotative approach, flaring as required to minimize forward speedat touchdown.

� Level aircraft. Increase collective pitch as contact is made with the water.

� When aircraft begins to roll, lower collective to full down to minimize bladesskipping off the water.

� Notify crew/passengers to evacuate aircraft after blades have stopped turning.

Do not inflate personal flotation gear until clear of the aircraft −safe exit will be restricted.

WATER LANDING − OEI/AEO

Conditions: Available power allows hovering.

NOTE: The gross weight of the aircraft will determine whether sufficient power isavailable to terminate the approach at a hover or whether a run−on landing mustbe performed if landing OEI.

Procedures:

� Establish normal approach to intended landing point.


� Plan to arrive at 100 FT above touchdown at approximately 40 KIAS.

� At approximately 50 FT, enter a decelerating attitude and increase power toreduce rate of closure.

� Descend to hovering altitude over water.

� Passengers and copilot exit aircraft.

WARNING






� Fly a safe distance away from all personnel in the water to avoid injury.

� Place ENGINE CONTROL switch(s) in OFF and perform a hovering autorota-tion.

� Allow aircraft to settle in a level attitude while applying full collective pitch.

� When aircraft begins to roll, reduce collective to full down to minimize bladesskipping off the water.

� Release safety harness and exit the aircraft as soon as the blades have stoppedturning.


Conditions: Available power does not allow hovering.

Procedures:

� Establish normal approach to intended landing point.


� Plan to arrive at 100 FT (30.5 M) above touchdown at approximately 40 KIAS.

� At approximately 50 FT (15.2 M), enter a decelerating attitude and increasepower to reduce rate of closure.

� As water contact is made, shut down engine and hold the helicopter as levelas possible, keeping forward speed and rate of descent to a minimum.

� When aircraft begins to roll, reduce collective to full down to minimize bladesskipping off the water.

� Release safety harness and exit the aircraft as soon as the blades have stoppedturning.


WARNING

WARNING




3−5. EEC MALFUNCTIONS

EECMANFAIL

EECMANFAIL

TORQUE

EEC MANUAL MODEANNUNCIATOR (YELLOW)

EEC CRITICAL FAULT WARNINGANNUNCIATOR (RED)

F92−021

EEC NON CRITICAL FAULTCAUTION ANNUNCIATOR

( YELLOW)

Figure 3−1. EEC Caution/Warning AnnunciatorsNOTE: The pilot should attempt to reset the EEC by by slightly moving the affected

engine’s twistgrip out of the NORMAL detent, pressing the EEC RESET button,and returning the twistgrip to NORMAL. Two attempts may be required. If theEEC malfunction indication clears, the EEC was experiencing a transient fault.If the EEC malfunction indication remains ON, the fault condition is still presentand the appropriate malfunction procedure shall be followed.

EEC CRITICAL FAULT

NOTE: In the event that the EEC on one of the engines fails, the fuel flow of that engineremains fixed and can only be controlled by the twistgrip. The engine with theserviceable EEC will attempt to maintain NP/NR within limits.

Indications: EEC with FAIL warning annunciator on and activation of the warning tonefor two cycles.

Procedures:

� When necessary, move the affected engine twist grip out of the NORMAL posi-tion to assume manual control of the FMU.

NOTE: The pilot has the option of leaving the fuel flow fixed or using the throttle twist gripto adjust the fuel flow (torque). When either twist grip is taken out of the NORMALposition, the EEC MAN annunciator will illuminate. Changes in power will becompensated through the serviceable EEC engine from zero torque totemperature limits. Twist grip movement is only required for large powerchanges.

When operating in manual mode (i.e., EEC MAN illuminated),reductions in power that allow the torque on the engine in theautomatic mode to approach zero % can lead to an increase inNP on the engine being manually controlled into the transient (20second time limit) overspeed range (> 104.5% − third yellow chevronis illuminated).

CAUTION






� Set power of the affected engine as desired.

� Continue flight and monitor engine indications on the IIDS primary display.

There is no NR governing following EEC failures on both engines.NR and power must be controlled by the pilot using a combinationof collective and twistgrips.

EEC NONCRITICAL FAULT

An EEC caution annunciator ON in flight may result in one of the following indica-tions.

Indications: Engine torque matching may be degraded.

Conditions: During flight

Procedures:

� Continue flight

� Advise maintenance

Indications: EGT indication blanks

Conditions: In flight

Procedures:

� Continue flight


Conditions: On ground prior to starting

Procedures:

� Do not attempt start, or abort start.


CAUTION




Indications: Inability to change engine mode with engine control switch.


Procedures:

� Continue flight

� After landing, perform manual engine shutdown (Ref. paragraph 3−7).

Conditions: On ground after landing

Procedures:

� Perform manual engine shutdown (Ref. paragraph 3−7).

Indications: NP and NR indications not matched (split)


Procedures:

� Continue flight

� Avoid maneuvers that cause NR to increase above normal. (High rates of de-scent, quick stops)

Conditions: On ground

Procedures:

� Do not takeoff







EEC MANUAL CONTROL

Indications: EEC and MAN annunciator ON

Procedures:

� EEC RESET switch PRESS

� Twistgrips CHECK IN NORMAL DETENT

Indications: EEC annunciator flashing. The EEC is in automatic, but one of the twist gripsis not in the normal position.

Procedures:

� Twist grip RETURN TO NORMAL POSITION

3−6. ENGINE STARTING − MANUAL

NOTE: The following procedure is provided to the pilot as a means of starting an engineafter experiencing an EEC FAIL warning which would preclude a normalautomatic start. Flight with one EEC failed (one engine manually controlled)should be considered an abnormal procedure. It should only be done toevacuate the helicopter from a hazardous environment or, if necessary, for areturn flight to a maintenance base where repairs can be performed. Beforeattempting a start and flight with an EEC FAIL warning on one engine, pilotsshould be familiar with the information in paragraphs 3−5 thru 3−7.

NOTE: To enable the starter to function during a manual start with an EEC FAIL warning,it may be necessary for a second crew member to push and hold the appropriateEngine Manual Start Button located on the back corners of the electrical loadcenter (Ref. Section VII).

NOTE: Complete the Engine Prestart cockpit check (Ref. Section IV) before attemptinga manual start.




� Collective control:

NOTE: The following steps of rotating the twist grip to reset the PLA are not required ifthe engine was previously shut down utilizing the manual shutdown proceduresin paragraph 3−7. Insure the twist grip is in the OFF position.

� � Twist grip on selected engine ROTATE TO FULL OPEN (PASTTHE ‘‘NORMAL’’ DETENT): THISRESETS THE PLA

� � Twist grip on selected engine ROTATE TO OFF

NOTE: AT a point between NORMAL and OFF, the twist grip will no longer be able tobe rotated toward the OFF position without applying additional force(approximately 30 to 40 lb / 13 to 18 kg).

� Electrical master panel:

� � Generator on selected engine OFF

� Fuel system panel:

� � L BOOST or R BOOST for appropriateengine

ON, CHECK IIDS INDICATIONS

� Engine control panel:

� � L ENGINE or R ENGINE IDLE

Monitor EGT, NG, and starter limits during start. Abort the startIf EGT rises rapidly through 700°C.CAUTION






� Abort start procedure:

� � Twistgrip on selected engine OFF

NOTE: AT a point between NORMAL and OFF, the twist grip will no longer able to berotated toward the OFF position without applying additional force (approximately30 to 40 LBS).

� � Engine control panel switch OFF WHEN EGT IS BELOW 150°C� Twist grip for selected engine ROTATE TOWARDS NORMAL

NOTE: As NG increases through 8% rotate twistgrip toward normal until lightoff occurs.Observe EGT indication for immediate temperature rise. Monitor EGT and NGduring start. Increase twistgrip toward normal only as necessary to keep NGaccelerating toward idle. Manually bring NP/NR to 65%.

If lightoff is not attained with an increase of EGT and NG within10 seconds, rotate the twistgrip to OFF and place the engine controlswitch to off. Following a 30 second fuel drain period, perform a30 second dry motoring run (Ref. Section VIII) before attemptinganother start. Repeat the complete starting sequence observinglimitations.

� Engine oil pressure CHECK

� Generator ON

� IIDS CHECK

� GPU start only:

� � Generators ON

� � GPU DISCONNECT

CAUTION




3−7. ENGINE/AIRCRAFT SHUTDOWN − MANUAL

NOTE: This procedure may be performed in the event a normal shutdown cannot beaccomplished on one or both engines.

� Collective stick FULL DOWN

� Cyclic stick TRIM TO NEUTRAL

� Pedals NEUTRAL

� Twist grip(s) IDLE DETENT

NOTE: The idle position is not marked on the twist grips. Idle is located at the point wherethe twist grip can no longer be rotated toward the OFF position without applyingadditional force (approximately 30 to 40 LBS).

� Utility panel:

� � All unnecessary bleed air and electricalequipment

OFF


� � L/R BOOST OFF


� � L/R GEN OFF

� NP(s) slows to idle CHECK

� EEC MAN indication(s) on primary IIDSdisplay

CHECK


� � L ENGINE or R ENGINE OFF

� Twistgrip(s) SNAP TO CUTOFF

NOTE: AT a point between NORMAL and OFF, the twist grip will no longer able to berotated toward the OFF position without applying additional force (approximately30 to 40 LBS).

� IIDS CHECK NORMAL SHUTDOWNINDICATIONS

� Continue with normal shutdown proce-dures






3−8. FIRE EMERGENCIES

CABIN FIRE/SMOKE

Indications: Smoke and fume accumulation in the cabin.


Procedures:

� Engine control switches OFF

� Passengers/crew EVACUATE

� Rotor brake (if installed) APPLY

� Power switch OFF


Procedures:

� Cabin heat OFF (if source of smoke is thecabin heat duct)

� Fresh air vents OPEN

� AC/VENT switch VENT LOW OR VENT HIGH

NOTE: If crew station and/or passenger compartment gaspers appear to be the sourceof smoke and or fumes, the AC/VENT switch should remain OFF or be returnedto OFF.

� Cockpit door vents OPEN

� Land immediately

� After landing:

� � Engine control switches OFF� � Rotor brake (if installed) APPLY� � Power switch OFF� � Passengers/crew EVACUATE




ELECTRICAL FIRE

Indications: Smoke and fume accumulation in the cabin.


Procedures:






Procedures:

� Cabin heat OFF


� Generator switches OFF

� If smoke/fire conditions persist:

� � POWER switch ESNTL� Land as soon as possible.

� After landing:







FIRECHIPS

FIRECHIPS

FIRECHIPS

LEFT ENGINE FIRE WARNINGANNUNCIATOR (RED)

TRANSMISSION DECK FIRE WARNING ANNUNCIATOR (RED)

RIGHT ENGINE FIRE WARNINGANNUNCIATOR (RED)

F92−022

Figure 3−2. Engine/Transmission Deck Fire Annunciators

ENGINE FIRE − ON GROUND

Indications: Engine FIRE warning annunciator ON and activation of the warning tone fortwo cycles.

Procedures:

� Engine control switches OFF BOTH ENGINES


� Fuel shutoff valve OFF FOR AFFECTED ENGINE

� Fuel boost pumps OFF

� Attempt to confirm existence of fire

� Fire bottle switch PRI (ALT IF NECESSARY)

NOTE: Fire bottle will not discharge with fuel valve ON.



ENGINE FIRE − DURING FLIGHT

Indications: Engine FIRE warning annunciator ON and activation of the warning tone fortwo cycles.

Procedures:

� Attempt to confirm existence of fire

� Engine control switch OFF FOR AFFECTED ENGINE




� Fuel shutoff valve OFF FOR AFFECTED ENGINE

� Airspeed REDUCE TO O.E.I. VNE OR LESS

� Fuel boost pump OFF FOR AFFECTED ENGINE

� Fire bottle discharge switch PRI (ALT IF NECESSARY)

NOTE: The fire bottle will not discharge with the fuel valve ON.

� If FIRE warning goes OFF LAND AS SOON AS PRACTICAL

� If FIRE warning remains ON LAND IMMEDIATELY

� After landing:

� � Engine control switch OFF FOR OPERATING ENGINE� � Rotor brake (if installed) APPLY� � Power switch OFF� � Passengers/crew EVACUATE

TRANSMISSION AREA FIRE

Indications: Transmission FIRE warning annunciator ON and activation of the warningtone for two cycles.

Procedures:


� After landing:

� � Engine control switches OFF� � Rotor brake (if installed) APPLY� � Power switch OFF� � Passengers/crew EVACUATE� � Fire extinguisher USE AS APPROPRIATE






3−9. FLIGHT CONTROL MALFUNCTIONS

ANTI−TORQUE FAILURE − GENERAL

Different types of failure may require slightly different techniques for optimumsuccess in recovery. Therefore, it is not possible to provide a standardized solu-tion for an anti−torque emergency.

ANTI−TORQUE FAILURE − COMPLETE LOSS OF THRUST

This involves a break in the fan drive system (ie., a broken drive shaft) that causesthe fan to stop turning resulting in a complete loss of fan thrust. Directional controlbecomes dependent on airspeed and power setting.

Indications: Inability to �trim" helicopter with pedals.

Conditions: In Forward Flight

Procedures:

� Adjust airspeed and power for level flight between 80 and 100 KIAS.

Do not attempt an autorotation from forward flight unless an actualdual engine failure occurs.

� Perform a shallow approach and running landing to a hard surface or othersuitable area. If possible, select an approach direction that offers a left quarter-ing headwind to reduce the touchdown ground speed and the amount of rightyaw.

NOTE: Touchdowns made into the wind between 20 and 30 KIAS, may provide gooddirectional control at reduced power (collective) settings.

� An aggressive reduction in power (collective) as the aircraft is deceleratedduring the final approach should yaw the aircraft to the left.

� As the ground is neared, adjust collective as necessary to align the air-craft with the touchdown direction and cushion the landing.

� During ground run out adjust collective to maintain directional control. If neces-sary, during touchdown and ground run out, reduce rotor RPM by rotatingboth twist grips simultaneously towards IDLE to assist in maintaining direc-tional control.

NOTE: Use of the twist grips to change RPM is generally not recommended due to thecomplexity of manipulating both twist grips simultaneously and now having bothengines in the manual mode. If needed, it is recommended that they be used onlyto reduce RPM just prior to or at the moment of touchdown.

CAUTION




Conditions: At a high hover

Indications: Helicopter begins an uncommanded turn to the right and does not respondto pilot input to the pedals.

Procedures:

� Reduce power with collective and attempt to fly away.

NOTE: If altitude permits, a positive reduction of collective pitch may result in a stoppingor slowing of the “uncommanded turn to the right”, and allow the pilot to fly outof the condition.

� Proceed with procedures for complete loss of thrust in forward flight

Conditions: At a low hover

Indications: Helicopter begins an uncommanded turn to the right and does not respondto pilot input to the pedals.

Procedures:

� Reduce power and altitude with collective, if necessary.

� As the ground is approached, rotate both twist grips simultaneously to IDLEand perform a hovering autorotation. Avoid rotating twistgrips with collectiveapplications during autorotation.

ANTI−TORQUE FAILURE − FIXED THRUSTER SETTING

Conditions: Right pedal applied

Procedures:

� Adjust airspeed and power for level flight at an airspeed that producesthe least amount of right yaw, usually between 80 and 100 KIAS.

� Perform a shallow approach and running landing to a hard surface or othersuitable area. If possible, select an approach direction that offers a left quarter-ing headwind to reduce the touchdown ground speed and the amount of rightyaw.

NOTE: Touchdowns made into the wind between 20 and 30 KIAS, may provide gooddirectional control at reduced power (collective) settings.

� An aggressive reduction in power (collective) as the aircraft is deceleratedduring the final approach should yaw the aircraft to the left.

� As the ground is neared, adjust collective as necessary to align the air-craft with the touchdown direction and cushion the landing.






� During ground run out adjust collective to maintain directional control. If neces-sary, during touchdown and ground run out, reduce rotor RPM by rotatingboth twist grips simultaneously towards IDLE to assist in maintaining direc-tional control.

NOTE: Use of the twist grips to change RPM is generally not recommended due to thecomplexity of manipulating both twist grips simultaneously and now having bothengines in the manual mode. If needed, it is recommended that they be used onlyto reduce RPM just prior to or at the moment of touchdown.

Conditions: Left pedal applied

Procedures:

� Use a shallow to normal approach into wind or with a right crosswind.

� Plan to touchdown with little or no forward speed.

� Maintain directional control with small adjustments in collective.

VSCS FAILURE

LEFT VERTICAL STABILIZERPOSITION INDICATOR

RIGHT VERTICAL STABILIZERPOSITION INDICATOR

MID−RANGE DEFLECTIONPOINT

L RVERTICAL STAB

L RVERTICAL STABFIN TRAILING EDGE

DEFLECTION INDICATORS

F92−023

Figure 3−3. VSCS Indicator

Indications: VSCS Fail message(s) on IIDS alpha−numeric display.

VSCS indicator: Abnormal indication − no movement or continuous full−scaledeflection.

Possible uncommanded sideslip in forward flight.

Procedures:

� Trim aircraft with pedals.

� VSCS OFF ON AFFECTED SYSTEM(S)

� Reduce airspeed below 100 KIAS (Ref. Section II).

� Continue flight to next point of intended landing.




HYDRAULIC SYSTEM FAILURE

Indications: Single system failure: The ‘‘1HYD’’ or ‘‘HYD2’’ caution annunciator illuminatedon the caution/warning advisory display. Both hydraulic system pressures or‘‘TEMPERATURE’’ indication will be displayed on the IIDS alphanumericdisplay. A stiffness in the anti−torque pedals will occur with a failure of thenumber 2 system.Dual system failure: The ‘‘1HYD2’’ caution annunciator illuminated on thecaution/warning advisory display. Both hydraulic system pressures or‘‘TEMPERATURE’’ indications will be displayed on the IIDS alphanumericdisplay.

Conditions: Single system failure − loss of pressure

Procedures:

� Decrease air speed to below 100 KIAS.

NOTE: A stiffness in the anti−torque pedals will occur with a failure of the number 2system.

� Continue the flight to the point of next intended landing.

� Perform a shallow approach to a hover; land vertically for a single systemfailure.

Conditions: Dual system failure − loss of pressure

Procedures:

� Decrease air speed to below 100 KIAS.

� Continue the flight to the point of next intended landing.

� Perform a running landing.

Conditions: High hydraulic fluid temperature

Procedures:

� Land as soon as practical.

CYCLIC TRIM FAILURE

Indications: Cyclic trim failure is indicated by an inability to reduce cyclic forces with thecyclic trim switch. Cyclic stick forces of approximately 15 lb (6.8 kg) may berequired for full control movement.

Procedures:

� Continue flight






COLLECTIVE FRICTION FAILURE

Conditions: Collective friction release failure

Indications: 25 lb (11.34 kg) of force required to move the collective up or down.

Procedures:

� Continue flight.

Conditions: Collective friction fails to engage

Indications: Collective control movements will require only 5 lb (2.27 kg) of force.

Procedures:

� Continue the flight.

3−10.PITOT/STATIC SYSTEM MALFUNCTION: SINGLE OR DUAL PITOT TUBE INSTALLATION

STATIC SYSTEM MALFUNCTION

Indications: Altimeter and IVSI (if installed) show no change in indication duringclimb/descent.

Conditions: Primary static source is clogged.

Procedures:

� Alternate static source toggle valve (on affected side)

PULL UP

NOTE: The altimeter will indicate 60 feet less during climb operations.

F92−024

CO−PILOT PITOT TUBE ALTERNATE STATIC SOURCETOGGLE VALVE LOCATED ON OPPOSITE SIDE OFINSTRUMENT PANEL.

NOTE: TO OPERATE TOGGLE VALVE, PULLVALVE HANDLE UP. TO RETURN TO PRIMARYSTATIC SOURCE, PUSH HANDLE DOWN.

Figure 3−4. Alternate Static Source Toggle Valve




PITOT HEAT FAILURE

Indications: Yellow pitot heat caution light (if installed) PITOT

L R

ON.

Conditions: Flight conditions requiring use of pitot heat.

Procedures: Left or right pitot heat failure.


Procedures: Left and right pitot heat failure.


3−11.ENGINE AND GENERATOR MALFUNCTION INDICATIONS

NOTE: Certain malfunctions may require an engine to be shutdown, however, the pilotmust assess the type of problem and decide if the affected engine is to remainoperational.

FIRECHIPS

%PSI°C

%LOAD

NG

GEN

ENGINE FIRE WARNINGANNUNCIATOR (RED)ENGINE CHIPS CAUTION

ANNUNCIATOR (YELLOW)

HIGH ENGINE OIL PRESSURECAUTION ANNUNCIATOR (YELLOW)

LOW ENGINE OIL PRESSURECAUTION ANNUNCIATOR (YELLOW)WARNING ANNUNCIATOR (RED)

HIGH ENGINE OIL TEMPERATUREWARNING ANNUNCIATOR (RED)

CAUTION ANNUNCIATOR(YELLOW)

ENGINE LOW OIL TEMPERATURECAUTION ANNUNCIATOR (YELLOW)

NG HIGH WARNING ANNUNCIATOR (RED)NG HIGH CAUTION ANNUNCIATOR (YELLOW)

GENERATOR OUT CAUTIONANNUNCIATOR (YELLOW)

GENERATOR HIGH LOADCAUTION ANNUNCIATOR (YELLOW)

GENERATOR LOADDIGITAL DISPLAY (WHITE)

F92−025

NG LOW WARNINGANNUNCIATOR (RED)

Figure 3−5. Engine and Generator Malfunction Annunciators






ENGINE HIGH OIL TEMPERATURE

Indications: Upper yellow caution annunciator ON at 120°C and/or Red warning annunciatorON at 125°C

NOTE: The engine is certified to operate continuously up to 125°C. The caution rangeand yellow annunciator are advisories only and indicate temperaturesapproaching maximum.

Procedures:

� Reduce power on affected engine.

� Monitor pressure and temperature.

NOTE: If temperature remains above limits (red annunciator ON) and/or abnormal oilpressure is indicated, shut down affected engine.

� If indications return to normal, increase power on affected engine as desired.


ENGINE LOW OIL TEMPERATURE

Indications: Lower yellow caution annunciator ON at +10°C and below for NG >50%.

Procedures:

� Allow engine oil temperature to increase to normal range before placing EngineControl in FLY.

ENGINE HIGH OIL PRESSURE

NOTE: The red high engine oil pressure annunciator is only activated during the lamptest mode.

Indications: Upper yellow caution annunciator ON if the oil pressure is greater than 100%PSI for more than 5 minutes.

Procedures:

� Monitor pressure.


� Advise maintenance.




ENGINE LOW OIL PRESSURE

Indications: Lower yellow caution annunciator or lower red warning annunciator ON.

Procedures:

� If single engine power is sufficient to continue flight, shut down affected engine.


ENGINE CHIPS

Indications: Yellow CHIPS caution annunciator ON.

Conditions: On ground:

� Shut down engine

Conditions: In flight:


NG HIGH

Indications: Red warning or yellow caution annunciator ON.

Procedures:

� Reduce power to normal range

� Check engine torque and EGT indications

NG LOW

Indications: Red warning annunciator ON.

Procedures:

� Check affected engine IIDS indications (primary display) for possible enginefailure.

� If engine failure is confirmed, proceed with engine failure procedures (Ref.paragraph 3−3).






GENERATOR HIGH LOAD

Indications: Upper yellow generator high load annunciator ON.

Procedures:

� Turn off unnecessary electrical equipment.

Failure to turn off unnecessary electrical equipment may causethe generator(s) to automatically go off line.

GENERATOR

Indications: Yellow GEN annunciator ON and %LOAD is ‘‘0’’.

Procedures:

� L GEN or R GEN (or both if dual generatorfailure) switch

RESET

� If GEN annunciator still ON OFF FOR AFFECTEDGENERATOR(S)


� If both generators failed.

� � Power switch ESNTL UNLESS FLIGHTCONDITIONS DICTATEOTHERWISE

NOTE: With both generators failed and the power switch in the ESNTL position, a fullycharged battery will supply power for at least 30 minutes.

WIth the power switch in the ESNTL position, only that equipment powered bythe essential bus will be operational.

� � Land as soon as practical.

CAUTION




3−12.TRANSMISSION MALFUNCTION INDICATIONS

FIRE

CHIPS

%PSI° C

TRANSMISSION CHIPS CAUTIONANNUNCIATOR (YELLOW)

TRANSMISSIONHIGH OIL PRESSURE WARNING ANNUNCIATOR (RED)

TRANSMISSIONHIGH OIL TEMP WARNING

ANNUNCIATOR (RED)

TRANSMISSIONHIGH OIL TEMP CAUTIONANNUNCIATOR (YELLOW)

TRANSMISSIONLOW OIL TEMP

CAUTIONANNUNCIATOR

(YELLOW)

TRANSMISSIONHIGH OIL PRESSURE CAUTIONANNUNCIATOR (YELLOW)

TRANSMISSIONLOW OIL PRESSURE CAUTION ANNUNCIATOR (YELLOW)

TRANSMISSIONLOW OIL PRESSUREWARNINGANNUNCIATOR (RED)

FIRE WARNINGANNUNCIATOR (RED)

F92−026

Figure 3−6. Transmission Malfunction Annunciators

TRANSMISSION OIL TEMPERATURE HIGH

Indications: Upper yellow caution or red annunciator ON.

Procedures:

� Reduce power

� Transmission oil pressure CHECK

� Air−conditioner (if installed) OFF

� If temperature remains high LAND AS SOON AS POSSIBLE

TRANSMISSION OIL TEMPERATURE LOW

Indications: Lower yellow caution annunciator ON.

Procedures:

� Continue flight

� Do not takeoff with low temperature annunciator ON.






TRANSMISSION OIL PRESSURE LOW

Indications: Lower red warning annunciator ON.

Conditions: Loss of transmission oil pressure.

Procedures:

� Reduce power to 56% torque or less as soon as possible.

� Land as soon as possible.

NOTE: The transmission has demonstrated operation without oil for 30 minutes at apower setting of 56%.

Indications: Lower yellow caution annunciator ON.

Procedures:


TRANSMISSION OIL PRESSURE HIGH

Indications: Upper yellow caution annunciator ON or red warning annunciator ON.

Procedures:

� Monitor transmission oil pressure.


TRANSMISSION CHIPS

Indications: Yellow CHIPS annunciator ON.

Procedures:

� Reduce airspeed to 100 KIAS.

� Monitor transmission oil temperature and pressure. If normal, land as soonas practical.

� If temperature/pressure are not normal, land as soon as possible.




TRANSMISSION INPUT TORQUE SPLIT

Indications: TQ SPLIT EXCEED message on the alphanumeric display.Possible EEC caution indicator on.

Conditions: Possible EEC noncritical fault resulting in transmission input torque mismatchof 18% or more.

NOTE: For this message to be displayed, both engines must be in the FLY mode, neitherengine can be out or operating in the manual mode, and the aircraft must not beon the ground as determined by AOG logic.

Procedures:








3−13.FUEL SYSTEM DISPLAY ADVISORIES

CAB HEATBAT HOT

BAT WRM

ROTORBRAKECABINDOOR

BAGGAGEDOOR

1 HYD 2

IIDS

FIRECHIPS

%LOAD

FUEL

GEN

LBOAT

°C

%PSI°C

FIRECHIPS

FIRECHIPS

NG

%PSI°C %PSI°C

%LOAD

NG

GEN

ÇÇÇÇÉÉÉÉ

FUEL

LB

ÇÇ YELLOWWHITE

B. NORMAL FUEL FLOW WITH IMPENDING ENGINE FUEL FILTERS BY−PASS

E. INDICATES BOTH FUEL SHUTOFF VALVES IN CLOSED POSITION (NOTE 1)

FUEL FILTER IMPENDING

BY−PASS CAUTION (YELLOW)

FUEL FILTER IMPENDING BY−PASS

CAUTION

FUEL SHUTOFF VALVE (YELLOW)

(NOTE 1)

FUEL SHUTOFFVALVE (YELLOW)

(NOTE 1)

ÉÉÂÂÂÂ

C. INDICATES LOW FUEL PRESSURE ON BOTH FUEL BOOST PUMPS OR BOTH BOOSTER PUMPS IN OFF POSITION

REDGREEN

LOW FUELWARNING

TICK MARK (RED)(LEFT SHOWN,

(RIGHT OPPOSITE) LOW FUEL CAUTIONSEGMENT (YELLOW)

FUEL QUANTITY SEGMENTS

FUEL QUANTITY DIGITALDISPLAY (WHITE)

D. LOW FUEL PRESSURE LEFT BOOST PUMP WITH NORMAL FUEL PRESSURE ON RIGHT BOOSTER PUMP

A. NORMAL FUEL PRESSURE

NOTE:

1. THE LIGHT SEGMENT BAR(S) WILL FLASH ON THE IIDS PANEL WHENTHE VALVE IS IN TRANSIT BETWEEN THE OPEN AND CLOSED POSITION.

LOW FUEL WARNING SEGMENT(RED)

F92−027

Figure 3−7. Fuel System Advisory Indicators




FUEL LOW

Indications: Fuel quantity displays yellow caution bars when fuel level decreases to 300pounds; the red warning bar displays at 150 pounds. Low fuel warningtick mark(s) displays at 127 to 97 LBS while in cruise flight.

NOTE: Under normal operating conditions (cruise flight), a low fuel warning tick mark(left or right) will illuminate when approximately 127 to 97 total pounds remainin the fuel tank. With tick mark(s) illuminated and both engines operating at MCP,approximately 10 minutes of fuel remain.Under conditions where either side of the fuel tank fails (i.e.,develops asubstantial leak) the system will display a low fuel warning tick mark whenapproximately 10 minutes of fuel remain (65 LBS) on either side of the collectortank at maximum OEI fuel consumption rate.

Procedures:

� With low fuel warning tick mark(s) ON 140 KIAS MAXIMUM AND AVOIDUNCOORDINATEDTURNS/MANEUVERS

Indications: Early display of low fuel warning tick marks − above 220 LBS in hoverand 160 LBS in cruise.

Conditions: Fuel transfer system malfunction.

Procedures:

� Place L BOOST and R BOOST switches OFF.

NOTE: Expect engine flameout on side with early low fuel warning tick mark illuminated.

FUEL BOOST PUMP FAILURE

Indications: Alternating white and yellow offset segments indicate low fuel pressure.

Procedures: Single Failure

� Place L BOOST and R BOOST switches OFF.

NOTE: If helicopter is equipped with the Supplemental fuel system, refer to Section X,“Operating Instructions: Supplemental Fuel System” for information regardingfuel transfer with boost pumps off.

For operation with Secondary Fuels (Ref. Section II), continue flight and avoidhigh ‘‘G’’ maneuvers.

� Continue flight

NOTE: If flight is continued into low fuel conditions (fuel warning tick mark(s) on), it ispossible for an engine to flame out from fuel starvation with as much as 50 LBSof fuel still being indicated on the fuel quantity display. Under this condition, theindicated fuel is available for OEI flight using the remaining engine.






SINGLE FUEL PROBE FAILURE

Indications: Digital fuel quantity indicator blanked.Vertical fuel quantity segments indicate approximately half the remaining fuelquantity.

Procedures:

NOTE: Continuous display of fuel flow is available on the IIDS as a top level menu item:

L ENG WF XXX PPH

L ENG WF XXX PPH

� Continue the flight using consumption and time calculations.

DUAL FUEL PROBE FAILURE

Indications: Digital fuel quantity indicator blanked.Vertical quantity segments blanked.

NOTE: The low fuel warning tick mark indication remains operational with a dual fuelprobe failure.

Procedures:

� Continue the flight using consumption and time calculations.

NOTE: Continuous display of fuel flow is available on the IIDS as a top level menu item:

L ENG WF XXX PPH

L ENG WF XXX PPH

IMPENDING FUEL FILTER BYPASS

Indications: Impending bypass is shown by an inverted ‘‘U’’ above affected fuel flow line.

Procedures:

� Continue flight

� If other bypass indicator is displayed LAND AS SOON AS POSSIBLE




FUEL SHUTOFF VALVE MALFUNCTION

Indications: Two yellow bar segments flashing above and below the fuel flow line to the left or right of center.

Conditions: Fuel valve not fully opened/closed

Procedures:

� In flight:

� � Continue flight

� � Be prepared for affected engine to flame out

� Pre Start:

� � Fuel shutoff switch CYCLE OFF TO ON

� � If no change in indication DO NOT ATTEMPT START






3−14.CAUTION AND WARNING ADVISORIES

CAB HEAT

BAT HOT

BAT WRM

ROTORBRAKECABINDOOR

BAGGAGEDOOR

1 HYD 2

IIDS

OAT°C

CABIN DOOR OPENCAUTION ANNUNCIATOR (YELLOW)

BAGGAGE DOOR OPENCAUTION ANNUNCIATOR (YELLOW)

IIDS MALFUNCTIONCAUTION ANNUNCIATOR (YELLOW)

HYDRAULIC SYSTEMPRESSURE OR HIGH TEMPERATURECAUTION ANNUNCIATOR (YELLOW)

CABIN HEATWARNING ANNUNCIATOR (RED)

BATTERY WARM CAUTION ANNUNCIATOR (YELLOW)

BATTERY HOTWARNING ANNUNCIATOR (RED)

ROTOR BRAKECAUTION ANNUNCIATOR (YELLOW)

F92−028

Figure 3−8. Caution/Warning Cluster

CABIN HEAT

Indications: Red CAB HEAT annunciator ON and activation of the warning tone for twoseconds.

Conditions: Bleed air leak

Procedures:

� Turn CAB HEAT switch OFF.




BATTERY HOT

Indications: Red BAT HOT warning annunciator ON (battery internal temperature 71°C)and activation of the warning tone for two cycles.


Procedures:

� Shut down aircraft.

� Service or replace battery prior to next flight.

Overheated battery can cause burns to personnel unlessprotective clothing and adequate tools are utilized. In someinstances the battery may cause a secondary fire or may ruptureadding the further danger of electrolyte burns. Exercise cautionin dealing with an overheated battery. Maintain extinguisher readyfor use. Do not use the fire extinguisher to cool the battery.


Procedures:

� Power switch OFF.

� Land as soon as possible

WARNING






BATTERY WARM

Indications: Yellow BAT WRM annunciator ON (battery internal temperature 57°C).


Do not attempt to start an engine on battery power with BAT WRMannunciator ON.

NOTE: A battery warm condition results in the battery being disconnected from theaircraft electrical system once a generator is placed on line. Generator poweralone is not sufficient to start an engine.

Procedures:

� Utilize a GPU to start engines.

� Power switch OFF after both generators are on line.

� If BATT WARM annunciator remains ON for more than five minutes, shutdownthe aircraft.

� Otherwise, continue flight.



Procedures:

� Power switch OFF.



CAUTION




ROTOR BRAKE

Indications: Yellow ROTOR BRAKE annunciator ON.

Procedures:

� Rotor brake handle CHECK STOWED

� If annunciator remains on, land as soon as possible.

CABIN DOOR

Indications: Yellow CABIN DOOR annunciator ON.


Procedures:

� Close and safe lock door

Conditions: In the air

Procedures:

� Reduce airspeed to 60 KIAS (Ref. Section II)

� Land as soon as practical and close and safe lock the door.

BAGGAGE DOOR

Indications: Yellow BAGGAGE DOOR annunciator ON.


Procedures:

� Close and safe lock door

Conditions: In the air

Procedures:

� Land as soon as practical and close and safe lock the door.






IIDS

Indications: Yellow IIDS annunciator ON.

Conditions: IIDS fault

Procedures:


� Check fault log after landing; advise maintenance.

3−15.OTHER MALFUNCTION/ADVISORIES

IIDS FAILURE

Indications: IIDS displays blanks.

Conditions: Loss of electrical power to IIDS.

Procedures: On ground

� Shut down.

Procedures: In flight

� Reduce airspeed to 100 KIAS or less.

� Reduce electrical load.


BATTERY DISCHARGE

Indications: BATT DISCHARGE message on IIDS alphanumeric display.

Conditions: Battery bus voltage is less than 26 volts

Procedures:

� Check generator load indications.

� Recycle GEN switches as required.




EXTINGUISHER PRESSURE LOW

Indications: EXTNGSHR PRESS LO message on IIDS alphanumeric display

Conditions: Low pressure in Halon containers.

Procedures: Advise maintenance

IPS BYPASS

Indications: IPS BYPASS message on IIDS alphanumeric display.

Conditions: Both IPS bypass doors open.

Procedures: Advisory only

NACA INLET MALFUNCTION

NOTE: Helicopters with the standard engine inlet screen do not have NACA doors.

Indications: NACA DOOR message on IIDS alphanumeric display.

Conditions: NACA door(s) in the incorrect position.

Procedures:

NOTE: In the event that the malfunction results in one door remaining closed after theaircraft has gone beyond the threshold airspeed of 47 KIAS, the engine with theclosed NACA door will indicate a higher EGT than the engine having the NACAdoor open.

� Place NACA INLET switch in CLOSE if flying in falling or blowing snow (Ref.Section II).

NOTE: Flight into falling or blowing snow is only permitted when the NACA inlet switchis in the closed position. The switch shall remain in the closed position for theduration of the flight, even after leaving the falling or blowing snow conditions.


ROTOR SPEED DISPLAY MALFUNCTION

Indications: Rotor speed display blanks.

Procedures: Avoid high rates of descent and maneuvers that would cause the rotor tooverspeed (e.g., rapid decelerations, quick stops, etc.)






GROUND POWER UNIT DOOR OPEN

NOTE: Helicopters with aft mounted batteries only.

Indications: Yellow GPU

indicator light ON.

Conditions: External power door open.

Procedures: On ground.

� Close door. If light remains ON with door closed, advise maintenance aftercompletion of flight.

� In the air: advise maintenance after completion of flight.

Procedures: In the air

� Advise maintenance after completion of flight.

3−16.VIBRATIONS

Indications: Sudden, unusual or excessive vibrations occurring during flight.

Conditions: The onset of unusual or excessive vibrations in the helicopter may be anindication of problems in the rotor or drive train systems.

Procedures:

� LAND AS SOON AS POSSIBLE.

� No further flights should be attempted until the cause of the vibrationhas been identified and corrected.

Indications: �CHECK NOTAR BAL" or �CHECK ROTOR BAL" on IIDS alphaneumericdisplay.

Conditions: NOTAR fan or main rotor balance out of acceptable range.

Procedures:

� Clear message from adphaneumeric display.


If the message on alphaneumeric display reappears during thesame flight, land as soon as possible.


CAUTION




3−17.EMERGENCY EGRESS

Crew compartment doors:Both doors function as primary and emergency exits.

Cabin door window removal:Each cabin door window may be used as an emergency exit by pulling the emergencyexit pull tab and pulling the window inward (Ref. Figure 3−9).

EMERGENCY EXIT

PULL TABTO REMOVE WINDOW

F92−029

EMERGENCY EXIT RELEASE

COCKPIT DOOR FRAMELOOKING OUTBOARD

RIGHT SIDE

RIGHT SHOWNLEFT OPPOSITE

1. LOCATED ON UPPER COCKPIT DOOR FRAME

2. LOCATED ADJACENT TO COCKPIT DOORACCESS HANDLE

CABIN DOOR EMERGENCY EXIT

CREW COMPARTMENT DOOR EXIT

Figure 3−9. Cabin Door Emergency Exit





FAA ApprovedReissue 1Original 3−43/(3−44 blank)

3−18.EMERGENCY EQUIPMENT

Emergency Fire Extinguisher:

The fire extinguisher mounts either to the aft side of the center console or onthe aft right hand side of the station 155.5 bulkhead. It detaches from the mount-ing bracket by unfastening the quick release clamps. The extinguisher uses Halon1211 extinguishing agent. The fire extinguisher is equipped with a pressuregauge that indicates normal, charge, and overcharge pressures.

F92−030

Figure 3−10. Emergency Fire ExtinguisherFirst Aid Kit:

The first aid kit is located on the right hand sidewall panel of the baggage compart-ment.


Normal Procedures

FAA ApprovedReissue 1Original 4−i/(4−ii blank)

S E C T I O N I VNORMAL PROCEDURES

TABLE OF CONTENTS

PARAGRAPH PAGE4−1. Preflight Requirements 4−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−1. Pilot’s Preflight Guide (Sheet 1 of 2) 4−2. . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−1. Pilot’s Preflight Guide (Sheet 2 of 2) 4−3. . . . . . . . . . . . . . . . . . . . . . . . . . .

4−2. Pilot’s Daily Preflight Check 4−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−3. Pilot’s Preflight Check 4−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−2. Instrument Panel − Single Pilot (Typical) 4−16. . . . . . . . . . . . . . . . . . . . .

Figure 4−3. Instrument Panel − Two Pilot (Typical) 4−17. . . . . . . . . . . . . . . . . . . . . . .

Figure 4−4. Switches and Circuit Breakers − Console Mounted (Typical) 4−18. . . .

Figure 4−5. Circuit Breakers − Baggage Compartment Mounted (Typical) 4−19. . .

Figure 4−6. Collective Pitch Stick Controls 4−20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−7. Cyclic Stick Grip 4−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−4. Engine Pre−Start Cockpit Check 4−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−5. Engine Starting − Automatic 4−24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−6. Engine Runup 4−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−7. Before Takeoff 4−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−8. NOrmal Takeoff 4−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−9. Cruise 4−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−10. Slow Flight/Approach 4−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−11. Landing 4−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−8. Tail Skid to Landing Surface Clearance 4−27. . . . . . . . . . . . . . . . . . . . . . .

4−12. Engine/Aircraft Shutdown − Normal 4−28. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−9. Cyclic Centering 4−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−13. Post Flight 4−30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−14. Noise Impact Reduction Procedures 4−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−15. Flight With Doors Removed or Cabin Doors Open 4−32. . . . . . . . . . . . . . . . . . . . . . . . .

Figure 4−10. Cabin Door Hold Open Device 4−32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−16. One Engine Inoperative Training 4−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

4−17. Fuel System 4−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


MD900 (902 Configuration with PW 207E)Normal Procedures


SECTION IVNORMAL PROCEDURES

4−1. PREFLIGHT REQUIREMENTS

NOTE: The checks described in this Section apply to the standard configuration MD900and do not include certain optional equipment items. Preflight checks for optionalequipment items may be found in Section X of this manual. If your helicopter isequipped with STC’d items, refer to the STC holder’s flight manual supplement.

‘‘CHECK’’ means to observe the helicopter and note any obvious damage. Damageis defined as any condition that is not normal or not within limits. Examples ofconditions to look for are: inoperable equipment, excessive leakage, discolorationcaused by heat, loose attachment, dents, cracks, punctures, abrasion, chafing, gall-ing, nicks, scratches, delamination and evidence of corrosion. These are the mostcommon types of damage, however, checks should not be limited to these items.

Further checks shall be performed before the next flight if discrepancies are notedto determine if the aircraft is airworthy. Flight is prohibited when unrepaired damageexists which makes the aircraft unairworthy.

Have a thorough understanding of operating limitations. (Ref. Section II).

Service helicopter as required. (Ref. Section VIII and the Aircraft MaintenanceManual).

Determine that helicopter loading is within limits. (Ref. Sections II and VI).

Check�helicopter�performance�data.�(Ref. Sections V, IX,�and�X).

Be sure to include a review of the appropriate flight manualsupplemental data for type of optional equipment installed(including STC items) as a regular part of preflight planning.

Perform Pilot’s Daily Preflight check prior to the first flight of the day.

Perform Pilot’s Preflight Check prior to subsequent flights that same day.

Brief passengers on relevant operational procedures and associated hazards (Ref.Section I).

CAUTION




F92−031−1A

Figure 4−1. Pilot’s Preflight Guide (Sheet 1 of 2)




F92−031−2A

Figure 4−1. Pilot’s Preflight Guide (Sheet 2 of 2)




4−2. PILOT’S DAILY PREFLIGHT CHECK

Perform these checks prior to the first flight of the day.

PRELIMINARY CHECKS

� Aircraft attitude CHECK

� Covers and tiedowns REMOVE

� Main rotor blades CHECK

EXTERIOR CHECKS − FRONT

� Battery compartment (front mounted bat-tery only)

BATTERY CONNECTED; SECURE

� Battery compartment door (front mountedbattery only)

CONDITION; CLOSED

� Pitot tube(s) CONDITION, FREE OFOBSTRUCTIONS

� Windscreen CONDITION

� Chin windscreen CONDITION

� � Chin windscreen area CHECK PEDAL LINKAGES;FOREIGN OBJECTS

� Fuselage Bottom:

� � Landing and searchlight SECURITY, CONDITION

� � Antennas/externally mounted equip-ment

CHECK SECURITY/ATTACHMENT

� � OAT probe CHECK




FORWARD RIGHT SIDE

� Right crew door:

� � Glass and vents SECURITY, CONDITION

� � Hinges CHECK

� � Latch system and handle CHECK OPERATION

� � Door release handle CONDITION

� � Door strut and strap OPERATION, SECURITY, CONDITION

� Crew Seat CONDITION; NOTHING STOWEDUNDER SEAT

� Pilot pedals:

� � Adjust ADJUSTMENT PINS ENGAGED

� � Move pedals by hand OBSERVE MOVEMENT OF THRUSTER


� � POWER switch BAT/EXT

� IIDS panel: CHECK LAMP TEST MODE (B.I.T.)

� � Fuel quantity CHECK

NOTE: The fuel quantity indication will not display actual fuel weight when the fuelsystem is ‘‘topped off’’. Remove fuel cap and pull lanyard to assure tank is fullby noting fuel level on the inside of filler neck (Ref. SECTION VIII). Fuel can betrapped in filler neck by the flapper valve.

� Lighting control panel CHECK OPERATION OF LIGHTS ASREQUIRED


� � POWER switch OFF

� Crew door CLOSE




FUSELAGE − RIGHT SIDE

� Landing gear:

� � Skid tube step SECURITY, CONDITION

� � Forward spacer fitting; crosstube CHECK

� � Skid tube and abrasion strips CHECK

� � Aft crosstube and damper fluid level CHECK (REF. SECTION VII)

� Aft fuel vent fairing CLEAR OF OBSTRUCTIONS



� Fuel sump drain:

� � Push in fuel drain control to take sample CHECK FOR CONTAMINATION;VERIFY PROPER OPERATION OFDRAIN VALVES

� � Fuel drain door CLOSED

� Forward fuselage skin and steps CONDITION

� External power door, (front mounted bat-tery only) avionics access panel, static port,fuel cap

SECURITY, CONDITION

� Right side passenger door:

� � Upper and lower track and guide CHECK

� � Door rollers CHECK OPERATION

� � Door skin and glass CHECK

� � Door stops/pins CHECK

� Right side passenger compartment:

� � Upholstery CHECK CONDITION

� � Seats and seat belts CHECK OPERATION

� � Cabin heat controls AS DESIRED

� � Loose equipment STOWED

NOTE: Nothing stowed under seats that are to be occupied.




RIGHT FORWARD TRANSMISSION DECK

� Hydraulic System:

� � Hydraulic manifold CHECK MOUNTING AND FLUID LEVEL� � System filters CHECK IMPENDING BYPASS

INDICATORS (REF. SECTION VIII)� � Longitudinal hydraulic actuator CHECK LEAKS, MOUNTING� � Hydraulic lines CHECK LEAKS, FITTINGS� Static mast supports CHECK ATTACHMENT� Rotor brake fluid level CHECK (IF INSTALLED)� Environmental control system:

� � Air inlet screen CHECK� � Cabin air and fan plenum CHECK MOUNTING� � Air ductwork CHECK CONDITION� Generator cooling ducts (if installed) CHECK CONDITION

� Transmission deck CHECK FOR FOREIGN OBJECTS ANDSIGNS OF FLUID LEAKAGE

� Forward access door CHECK OPERATION AND CONDITION;CLOSE

� Generator cooling inlet (if installed) CHECK

RIGHT CENTER TRANSMISSION DECK

� Oil cooler:

� � Cooling air inlet NO OBSTRUCTIONS

� � Oil cooler CHECK MOUNTING, LEAKS ANDCONDITION

� � Air ducts CHECK MOUNTING AND CONDITION

� Transmission:

� � Transmission oil level CHECK

� � Transmission oil filler cap CHECK SECURITY

� Static mast support CHECK MOUNTING AND CONDITION

� Forward outside engine mount CHECK MOUNTING AND CONDITION




� Engine drive shaft CHECK

� Fan drive shaft CHECK

� Rotor brake CHECK


� Engine accessory gear box CHECK FITTINGS, LINES, CONNECTORSAND WIRING

� Engine oil filter CHECK BYPASS INDICATOR

� Transmission access door latches,hinges, and door

CHECK OPERATION; CLOSE

� Cabin door closed and latched CHECK

FUSELAGE − RIGHT TOP REAR

� Work platforms/steps. CHECK

� Engine air inlet w/o particle separator:

� � Inlet screen CHECK − NO OBSTRUCTIONS

� � NACA inlet NO OBSTRUCTIONS

� Engine air inlet with particle separator:

� � Particle separator CHECK − NO OBSTRUCTIONS

� � Bypass door CLOSED − CONDITION OF SEAL

� � NACA inlet door CLOSED − NO OBSTRUCTIONS

� Right Engine:

� � Engine oil access door CHECK CONDITION

� � Engine oil level CHECK

NOTE: To reduce the possibility of over servicing and ensure accurate readings for oilconsumption measurement, it is recommended that oil level always be checkedwithin 10 minutes after engine shutdown (Ref Section VIII).

� � Oil filler cap CHECK

� � Engine cowling assembly CHECK




� Fuselage skin CHECK CONDITION

� Notar fan inlet:

� � Fan air inlet screen and duct CHECK CLEAR

� � Notar fan blades CHECK

ROTOR SYSTEM

� Stationary swashplate CHECK

� Lower control rodend bearings CHECK

� Rotating swashplate CHECK

� Scissors drive link CHECK

� Pitch change links CHECK

� Striker plates and rollers CHECK

� Inner flexbeam attach points CHECK

� Flexbeam lead and lag legs CHECK

� Upper and lower damper and damper caps CHECK

� Elastomeric feathering bearing CHECK

� Pitch change housing CHECK

� Blade attach pins (bolts):

� � Check for upward shift of installed bladeretention bolts.

ADVISE MAINTENANCE IF UPWARDSHIFT IS NOTED

� � Check blade retention bolts for gap be-tween thrust washer and retainer.

ADVISE MAINTENANCE IF NO GAP ISPRESENT

� Blade attach points CHECK

� Rotor blades CHECK

� Top of rotor head CHECK




FUSELAGE − RIGHT REAR

� Fuselage skin CHECK

� Exhaust ejector cowl CHECK

� Baggage door:

� � Handle OPERATION

� � Skin CHECK

� � Door strut CHECK

� � Rear spoiler CHECK

� � Hinge pins CHECK

� � Environmental control system vent CHECK

� Baggage compartment:

� � Loose items SECURED

� � Circuit breaker panel CHECK

� � Baggage door CLOSED AND LATCHED



TAILBOOM AND EMPENNAGE − RIGHT SIDE

� Tailboom attach ring CHECK

� Tailboom slots CLEAR OF OBSTRUCTIONS

� Tailboom CHECK CONDITION

� Horizontal stabilizer:

� � Horizontal stabilizer attach points CHECK

� � Horizontal stabilizer fairing CHECK

� � Antennas CHECK SECURITY/ATTACHMENT

� Vertical stabilizer CHECK

� Nav light/strobe lenses CHECK




� Thruster rotating cone: CHECK FOR FREEDOM OF ROTATION

� � Place hands at the 11 and 5 o’clock positions and press inward while rotating thecone to the left and right. Repeat check by using the 1 and 7 o’clock positions. Ad-vise maintenance if any unusual noise or roughness is noticed.

CAUTION: Do not rotate cone beyond one−half left/right open.

� � Turning vanes CHECK

TAILBOOM AND EMPENNAGE − LEFT SIDE

� Horizontal stabilizer:

� � Horizontal stabilizer attach points CHECK

� � Horizontal stabilizer fairing CHECK

� � Antennas CHECK SECURITY/ATTACHMENT

� Tail skid CHECK

� Vertical stabilizer CHECK

� Nav light CHECK

� Tailboom CHECK CONDITION

� Tailboom attach ring CHECK

LEFT REAR FUSELAGE

� Fuselage skin CHECK

� Exhaust ejector cowl CHECK

� Work platforms/steps CHECK

� External power door, (aft mounted batteryonly)

SECURITY, CONDITION






FUSELAGE − LEFT SIDE

� Landing gear: CHECK

� � Aft crosstube and damper fluid level CHECK (REF. SECTION VII)

� � Passenger step CHECK

� � Skid tube and abrasion strips CHECK

� � Forward spacer fitting; crosstube CHECK

� � Skid tube step SECURITY, CONDITION

� Underside of fuselage:

� � Fuselage skin CHECK



� Left side passenger door:

� � Upper and lower track and guide CHECK

� � Door rollers CHECK OPERATION

� � Door skin and glass CHECK

� � Door stops/pins CHECK

� Left side passenger compartment:

� � Upholstery CHECK CONDITION

� � Seats and seat belts CHECK OPERATION

� � Loose equipment STOWED


FORWARD LEFT SIDE

� Left crew/passenger door:

� � Glass and vents SECURITY, CONDITION

� � Hinges CHECK

� � Latch system and handle CHECK OPERATION

� � Door release handle CONDITION

� � Door strut and strap OPERATION, SECURITY, CONDITION




� Crew Seat CONDITION − NOTHING STOWEDUNDER SEAT


� Co−pilot pedals (if installed) ADJUSTED; ADJUSTMENT PINSENGAGED

� Crew door CLOSE

� Avionics access panel CHECK

� Static port CHECK − NO OBSTRUCTIONS

� Fwd fuel vent fairing CHECK − NO OBSTRUCTIONS

LEFT FORWARD TRANSMISSION DECK

� Hydraulic System:

� � Hydraulic manifold CHECK MOUNTING AND FLUID LEVEL� � System filters CHECK IMPENDING BYPASS

INDICATORS (REF. SECTION VIII)� � Lateral and collective hydraulic actua-

torsCHECK LEAKS, MOUNTING

� � Hydraulic hand pump CHECK LEAKS, FITTINGS� � Hydraulic lines CHECK LEAKS, FITTINGS� Static mast supports CHECK ATTACHMENT� Environmental control system: (if installed)

� � Evaporator CHECK� � Freon lines CHECK� � Air ductwork CHECK CONDITION� Transmission deck CHECK FOR FOREIGN OBJECTS AND

SIGNS OF FLUID LEAKAGE� Transmission oil filter CHECK BYPASS INDICATOR (REF.

SECTION VIII)� Generator cooling ducts (if installed) CHECK CONDITION

� Forward access door CHECK OPERATION AND CONDITION;CLOSE

� Generator cooling inlet (if installed) CHECK




LEFT CENTER TRANSMISSION DECK

� Oil cooler:

� � Cooling air inlet NO OBSTRUCTIONS

� � Oil cooler CHECK MOUNTING, LEAKS AND CONDITION

� � Air ducts CHECK MOUNTING AND CONDITION

� Static mast support CHECK MOUNTING AND CONDITION

� Forward outside engine mount CHECK MOUNTING AND CONDITION

� Engine drive shaft CHECK

� Fan drive shaft CHECK

� Rotor brake CHECK


� Engine accessory gear box CHECK FITTINGS, LINES,CONNECTORS AND WIRING

� Engine oil filter CHECK BYPASS INDICATOR

� Transmission access door latches, hinges,and door

CHECK OPERATION; CLOSE

� Cabin door closed/open and latched/secured CHECK

FUSELAGE − LEFT TOP REAR

� Engine air inlet w/o particle separator:

� � Inlet screen CHECK − NO OBSTRUCTIONS� � NACA inlet NO OBSTRUCTIONS

� Engine air inlet with particle separator:

� � Particle separator CHECK − NO OBSTRUCTIONS

� � Bypass door CLOSED, CONDITION OF SEAL� � NACA inlet door CLOSED, NO OBSTRUCTIONS




� Left Engine:

� � Engine oil access door CHECK

� � Engine oil level CHECK

NOTE: To reduce the possibility of over servicing and ensure accurate readings for oilconsumption measurement, it is recommended that oil level always be checkedwithin 10 minutes after engine shutdown (Ref Section VIII).

� � Oil filler cap CHECK

� � Engine cowling assembly CHECK

4−3. PILOT’S PREFLIGHT CHECK

Perform these checks prior subsequent flights of the same day.

� Fluid levels CHECK

� Transmission deck − signs of fluid leakage CHECK

� Air inlet screens/particle separators CHECK

� Fuel cap, access doors and panels CHECK

� Rotor blades CHECK BLADE RETENTION BOLTS(PINS)

� Rotor blades CHECK

� Tailboom and empennage CHECK

� Cargo and loose equipment CHECK

� Baggage, cabin and crew doors CHECK




F92−032−1A

IIDS

AIRSPEEDINDICATOR

MAGNETIC COMPASS

ALTIMETER

IVSI

FOOT HEATERCONTROL VALVE

ALTERNATE STATIC SOURCETOGGLE VALVE

VSCS INDICATOR

ATTITUDEINDICATOR

TAKEOFF TIMINGINDICATOR LIGHTS

GPU DOORLIGHT (AFT MOUNTED

BATTERY ONLY) LOCATION TYPICAL

Figure 4−2. Instrument Panel − Single Pilot (Typical)




F92−032−2A

ALT

ER

NAT

E S

TATIC

SO

UR

CE

T

OG

GLE

VALV

EA

LTE

RN

ATE

STAT

IC S

OU

RC

E

TO

GG

LE VA

LVE

FO

R LE

FT

SID

E

NO

TE

1:N

OT

E 2:

NO

TE

1:N

OT

E 2:

PIT

OT

/STAT

IC IN

ST

RU

ME

NT

S.

Figure 4−3. Instrument Panel − Two Pilot (Typical)




NAVCOM 1

ESSBUS R

GCUR

ENGFIRE R

XMSNFIRE

FUELVLV R

PITOTHEAT R

IIDSVSCS

R TRIM AUDIOAP/SASALERT

INSTRLTG

EADIR PRIMARY

NAVCOM 2

LDGLT

AIRDATA CFU

AHRS 1

ESSBUS L

GCUL

ENGFIRE L

FUELVLV RL

VSCSL

R ESS BUS

L ESS BUSFUEL

PROBE GPS

BLD AIRLEAK

STBYATT

AP/SASEHSIR CMPTR DISC ACCEL

L DC BUS R DC BUSINSTRFLOOD ANNUN

XPNDR1

BLD AIRHEAT

COCKPIT HEAT

CONTROL

ENGINE CONTROLOVSP TEST

L ENGINE R ENGINE

OFF

IDLE FLYTRAIN

FUEL SYSTEML BOOST R BOOST

ON

OFF

BOTTLE

PRI

ALT

LEFT OFF RIGHT OFF

ON

OFF

FUEL SHUTOFF

DISCHARGE

OFF

ELECTRICAL MASTER

AVIONICS L GEN R GEN POWER

ON

OFF RESET RESET

ON ON

OFF OFF OFF

BAT/EXT

ESNTL

LIGHTING CONTROL

LT MSTR CONSOLE IIDS

FLOOD INSTR

STROBE AREAPOSN

ON

OFF OFF OFF

ON

OFF

ON

OFF OFF

BOTH

CKP

CAB

IPS

HEAT

AC/VENT

PITOTHEAT

HYDTEST

OFF

ON

SYS 1 OVRD

SYS 2

COOLLOW

COOLHIGH VENT

HIGH

VENTLOW

OFF

OFF

OFF

ON

ON

L VSCS RON

OFF

TEST

CAB

F927−006A

KEY SWITCHOFF

IDLE FLYTRAIN

UTILITY PANEL

NOTE

HELICOPTERS WITHIPS ONLY

STROBES

BOTH

RED

OFF

NACAINLET

CLOSE

NORMAL

Figure 4−4. Switches and Circuit Breakers − Console Mounted (Typical)




F927−007A

PITOTHEAT 2

LEFT GENERATOR BUS

AUDIOPNL 2

CKPTUTL

CABUTL

EVAP VENTEVAP

COMP

ATTGYRO 2

CPLTCLOCK

CNDSRFAN 2

EADIL

EHSIL

L W/SWIPER

AHRS1 AUX

LH DCFDR

IIDS TRAKSTB

HYDTEST

AVFAN

IPS

HOIST CUT

HOISTPWR

ATTGYRO1

PILOTCLOCK

CNDSRFAN 1

ELT R W/SWIIPER

AHRS2 AUX

RH DCFDR

FD SYN FLT DIR MODE SEL INVTRLEFT ESS BUS

LEFT AVIONICS BUS

RIGHT GENERATOR BUS

ADF2 RADARRT

RADARIND

MKRBCN

RADALT

PAPWR

COM 3 XPNDR2

DIRGYRO 2

NAV 3

MVGMAP

LIGHTING

RIGHT AVIONICS BUS

L R L R

L R L R

BST PUMP EECRH FUEL

LOW

DETENT IGNTR

CNSL POSN STROBE AREA

AHRS2 PRI

AVMSTR

AUXFUEL

FIREHRD

SMOKEDET

ENCALT

SRCHLGT

HVRLGT

NACA LH FUELFUEL

CABAUD

5 VDIM

NSUNCONT

NSUNPWR

CARGOHOOK

L FLDEXCIT

R FLDEXCIT

HDG SAS/AP ADF26 VAC BUS

ADF1 FMCTRL

FM1RT

FM2RT

FM3RT

DME STORMSCOPE

CAMERA NAV 1 RMI

BATTERY BUS

20

T/OTIMER

Figure 4−5. Circuit Breakers − Baggage Compartment Mounted (Typical)




45

6

7

8

9

10

1. COLLECTIVE FRICTION RELEASE2. EEC RESET SWITCH3. TAKEOFF TIMER4. HOVER, LANDING AND SEARCHLIGHT SWITCHES5. SEARCH LIGHT CONTROL SWITCH 6. GO−AROUND SELECT SWITCH7. COMMUNICATIONS SELECT SWITCH8. YAW SYNCHRONIZATION SWITCH9. AUTO PILOT YAW/VERTICAL BEEP SWITCH10. LEFT/RIGHT ENGINE TWIST GRIPS11. INDEX MARKS12. ALIGNMENT MARK

F927008

L

R

NORMAL

NORMAL

1

11

12

3

2

Figure 4−6. Collective Pitch Stick Controls




CYCLIC TRIM

FLOAT INFLATION SWITCH

RADIO/ICS

F92−036

AUTO PILOTDISENGAGE

CARGO HOOKRELEASE

Figure 4−7. Cyclic Stick Grip




4−4. ENGINE PRE−START COCKPIT CHECK

ELECTRICAL POWER − OFF

� All cabin doors closed and safelocked CHECK� Seat belt and shoulder harness for proper fit and engage-

ment of buckleFASTENED

� Operation�of�shoulder�harness�inertia lock CHECK� Rotor brake STOWED� Magnetic compass CHECK� Flight instruments CHECK STATIC

POSITION/SET� Collective Control:

� � Collective friction ON

� � Collective stick position FULL DOWN

If collective is not full down, do not try to force down until hydraulicpressure increases during start.

� � Twistgrip alignment marks aligned with index mark CHECK

� � LDG/HVR lights OFF� Key switch ON� Essential bus panels:

� � Circuit breakers IN� NACA inlet panel (if installed):

� � NACA inlet switch NORMAL� Utility panel:

� � CAB HEAT OFF

� � AC/VENT OFF

� � PITOT HEAT (if installed) OFF

� � IPS (if installed) OFF

� � VSCS L/R ON

CAUTION




� Lighting control panel:

� � LT MSTR AS REQUIRED

� � CONSOLE/IIDS/FLOOD/INSTR AS DESIRED

� � STROBE(S)/POSN/AREA AS DESIRED

NOTE: If white strobe lights are installed, the “BOTH” position is to be used duringdaytime operations only. (Ref. Figure 4−4.)


� � Avionics AS DESIRED

� � L GEN and R GEN ON (OFF FOR GPUSTART)

� � POWER OFF� Fuel system panel:

� � L BOOST AND R BOOST OFF

� � LEFT/RIGHT FUEL SHUTOFF ON; COVER CLOSED� Engine control panel:

� � L ENGINE and R ENGINE OFF

ELECTRICAL POWER − ON


� � POWER BAT/EXT

NOTE: If helicopter has the aft battery option, the yellow GPU light will be ON when aGPU is used for electrical power.

� IIDS:

� � Monitor BIT FIRE WARNING ANNUNCIATORSON FOR 2 SECONDS; CHECK IIDSFOR ADVISORIES

NOTE: Perform a commanded IIDS BIT if the helicopter has been statically exposed totemperatures below 0°C for 12 hours or longer.

� � Fuel quantity display CHECK

� � DISP (display by exception) AS DESIRED




4−5. ENGINE STARTING − AUTOMATIC

NOTE: Either engine may be started first.Engine starts have been demonstrated at temperatures as low as −36°C with aground power unit (GPU) assisted by the aircraft battery.Engine starts using battery power only have been demonstrated after the aircraftand battery have been statically exposed to temperatures down to 0°C for 12hours or more.A GPU should be used in lieu of aircraft battery power when attempting morethan one initial engine start during operations in ambient temperatures above32°C.Maximum wind speed for starting and stopping the rotor is 50 knots.


� � L BOOST or R BOOST ON; CHECK IIDS INDICATION

� EEC MAN indicators OFF


� � L ENGINE or R ENGINE SET TO IDLE/FLY AS REQUIRED

Monitor EGT, NG, and starter limits during start. If EGT is observedapproaching maximum overtemperature limits during start (Ref.Section II), abort the start as follows.Engine control switch OFF, fuel boost pump OFF; monitor IIDSdisplays.

If lightoff is not attained with an increase of EGT and NG within10 seconds, turn fuel boost pump OFF and place the engine controlswitch to OFF. Following a 30 second fuel drain period, performa 30 second dry motoring run (Ref. Section VIII) before attemptinganother start. Repeat the complete starting sequence observinglimitations. This procedure applies to ground and air−starts in theauto mode.

Abort start if: abnormal noises are heard; engine start hangs (NGbelow 54%); NG or NP increase beyond limits; start is not completedwithin 45 seconds.

Ensure collective full down, cyclic (Ref. Figure 4−9) and pedalscentered as hydraulic pressure increases. Should an abnormalvibration occur as the NR passes through 35 to 40%, shutdownaircraft and advise maintenance. This vibration may indicate thatpossible damage to the flexbeam has occurred.

If collective is not full down, do not try to force down until hydraulicpressure increases during start. Sufficient hydraulic pressure willbe available when NR is above 25 percent.

CAUTION




� IIDS CHECK FOR NORMAL INDICATIONS

NOTE: If any abnormal indications are observed, i.e. low transmission/engine oilpressures, shut down engine.

� Repeat starting procedure for second engine

NOTE: Do not start second engine until at least 60% NR is attained on the first engine.

� GPU start only:

� � L GEN/R GEN ON

� � GPU DISCONNECT

� � Yellow GPU indicator light (aft mounted battery only) OUT

4−6. ENGINE RUNUP


� � Avionics ON, AS DESIRED


� � L ENGINE and R ENGINE FLY

4−7. BEFORE TAKEOFF

� Cyclic response check:

� � Move cyclic stick and observe rotor tip for correct movement.

� Collective friction AS DESIRED

� Primary and secondary IIDS displays CHECK ADVISORIES

� Utility Panel:

� � PITOT HEAT (if installed) AS REQUIRED

NOTE: Turn pitot heat ON when visible moisture conditions prevail and OAT is 5°C andbelow.

� � IPS switch (if installed) AS DESIRED

� � CAB HEAT AS DESIRED




4−8. NORMAL TAKEOFF

� Hover area and takeoff path CLEAR

� Hover power NOTE TORQUE

� Takeoff PERFORM, USING UP TO 10% ABOVEHOVER POWER

NOTE: For takeoff in noise−sensitive areas, refer to Paragraph 4−14.

NOTE: With the fuel system ‘‘topped off’’, the IIDS fuel quantity will not display adecrease until after approximately 10 minutes of flight.

4−9. CRUISE

IPS switch (if Inlet Particle Separator installed) may be turned OFF.

NOTE: Decision to use the inlet particle separator scavenge air should be based onatmospheric conditions, gross weight and height above terrain where operationsare to be conducted.

NACA doors (if installed) may be closed if blowing dust, sand, etc. is present inthe atmosphere.

IIDS menu/display mode selection:

The IIDS MENU and DISP keys may occasionally become non−responsive and/ormay register key inputs twice. This condition will clear itself after a brief interval.

Care should be taken when using the arrow keys (‘‘�’’ or ‘‘�’’) to scroll betweenmenu and submenu names, or between data and message items. Pressing thearrow keys too fast may result in scrolling past the desired menu or messagedisplayed. The arrow keys should be pressed only after the menu item or messagechanges in the alphanumeric display.

If a garbled message appears while scrolling through a menu, scrolling pastthat menu item and then returning to the desired menu item will correct thedisplay.

While switching between the display modes (display by exception or continuousdisplay modes), pilots are reminded to clear the alphaneumeric display priorto switching display modes.

4−10.SLOW FLIGHT/APPROACH

Observe controllability envelope and critical wind azimuth as stated in Section II.

The NACA door actuators (if installed) receive a discrete input from an airspeedswitch in the airspeed indicator. This signals the NACA doors to automatically close.When airspeed increases above 47 KIAS, the NACA doors open. If door actuatorfails to function properly, the IIDS will display ‘‘NACA DOOR’’ advisory messagein the alphanumeric display.




4−11.LANDING

Use the illustration below to determine safe landing attitudes. Nose up attitudesin excess of 9° 40′ will result in the tail skid contacting the landing surface.

F927−098

9° 40′30.16 IN(76.61cm)

Figure 4−8. Tail Skid to Landing Surface ClearanceRunning landing:

Maximum recommended ground contact speed is 30 knots for smooth hard sur-face.

Avoid�rapid�lowering�of the�collective and aft cyclic after ground contact.

Slope landing:

Slope landings have been demonstrated up to 12° in any direction. Successfulcompletion of this maneuver on a particular surface will depend on sufficientfriction between the skid tubes and the landing surface to prevent the helicopterfrom sliding.




4−12.ENGINE/AIRCRAFT SHUTDOWN − NORMAL

NOTE: Shut down the engines before exiting the helicopter unless safety or operationalconsiderations dictate otherwise.

Maximum demonstrated wind speed for starting and stopping the rotor is 50knots.

� Collective stick FULL DOWN; FRICTION ON

� Cyclic stick TRIM TO NEUTRAL(REF FIGURE 4−9)

� Pedals NEUTRAL

� Engine control panel

� � L ENGINE and R ENGINE SET TO IDLE

� All unnecessary electrical equipment OFF

� Utility panel:

� � Heat OFF

� � AC (if installed) OFF

� � Pitot heat (if installed) OFF

� � IPS (if installed) OFF

� Lighting control panel AS DESIRED


� � Avionics master switch OFF

� � L GEN/R GEN switches OFF


� � L BOOST/R BOOST OFF

Failure to turn OFF boost pumps will result in engine fuel nozzlecoking over time.CAUTION




CAUTION: CYCLIC SHOULD BE TRIMMED TO THE NEUTRAL POSITION FOR START−UP AND SHUTDOWN.

NEUTRAL POSITION IS ACHIEVED WITH CENTERING STRAP EXTENDED, TOUCHING CENTERING DECAL WHEN PERPENDICULAR TO INSTRUMENT PANEL.

CENTERING STRAP INSTOWED POSITION)

CENTERING STRAP IN EXTENDED POSITION

CENTERING DECAL

CENTERING STRAP

90°

F92−037

Figure 4−9. Cyclic Centering


� � L ENGINE and R ENGINE OFF

� ENG OUT indications CHECK IIDS

Do not use collective pitch to slow rotor.Should an abnormal vibration occur as the NR passes through 40to 35%, advise maintenance before further flights. This vibrationmay indicate that possible damage to the flexbeam has occurred.

NOTE: Check that compressor decelerates freely. Abnormal noise or rapid run down(rapid loss of NG) may indicate turbine blade rubbing.

If there is evidence of post engine high EGT, follow the dry runprocedure as described below.

� Dry run procedure:

� � Twist grip OFF� � Engine control switch for selected engine SET TO IDLE −

OBSERVE STARTERTIME LIMITS

� � Engine control switch for selected engine OFF

CAUTION

WARNING




Normal shutdown continued:

� Rotor brake (if installed):

� � Raise brake handle to release from stowed position

� � Rotate handle clockwise and apply brake by pullingdown on handle until handle locks aft. Release rotorbrake during last revolution unless conditions dictateotherwise.

APPLY BELOW70% NR

Care should be taken while applying the rotor brake if the helicopteris parked on a slippery or icy surface. Anti−torque control isminimized at less than normal operating RPM when the engine isnot driving the rotor system. Full control of the helicopter duringthese conditions may be limited.

� IIDS CHECK FORINDICATIONS ORMESSAGES

NOTE: If entering the IIDS “Time Summary” menu to check “TOT FLT HR”, turn powerswitch to OFF after 0% NG. then back to on to check “TOT FLT HR”.


� � POWER OFF AT 0% NG� Key Switch AS DESIRED

4−13. POST FLIGHT

� Aircraft−investigate�any�suspected damage CHECK

� Rotor blades CHECK BLADERETENTION BOLTS(PINS)

� Fuel and oil leaks CHECK

� Engine and rotor transmission oil levels CHECK

NOTE: Engine oil level should be checked within 10 minutes after shutdown.

� Logbook entries COMPLETE

� Flight manual and equipment STOWED

� Aircraft tiedowns, covers AS REQUIRED

CAUTION




4−14.NOISE IMPACT REDUCTION PROCEDURES

Safe operation of the helicopter always has the highest priority.Utilize the following procedures only when they will not conflictwith safe helicopter operation.

Certain flight procedures are recommended to minimize noise impact on surroundingareas. It is imperative that every pilot subject the public to the least possible noisewhile flying the helicopter.

Takeoff:

Takeoff using maximum takeoff power at the speed for best rate of climb (Ref.Section V).

Proceed away from noise sensitive areas.

If takeoff must be made over noise sensitive area, distance (altitude) is the bestform of noise suppression.

Cruise:

Maintain 1000 feet minimum altitude where possible.

Maintain speed of no more than 110 KIAS over populated areas.

Coordinated turns at around the speed for best rate of climb cause no appreciablechange in noise.

Sharper turns reduce area exposed to noise.

Approach:

Use steepest glideslope consistent with passenger comfort and safety.

Noise characteristics data is provided in Section V.

CAUTION




4−15.FLIGHT WITH DOORS REMOVED OR CABIN DOORS OPEN

Stow or secure all loose objects with doors opened or removed.

The aircraft may be flown with cabin doors open or removed in accordance withthe flight restrictions stated in Section II.

NOTE: Refer to Section VI for weight and balance data with doors opened or removed.

One or both cabin doors may be opened or closed in flight at airspeeds upto 60 KIAS.

For sustained flight with the cabin doors open, use of the cabin door holdopen device is required (Ref. Figure 4−10).

CABIN DOORRESTRAINT FITTING

LEFT SIDE, LOOKING INBOARD

F92−038

NOTE: THE CABIN DOOR HOLD OPEN DEVICE OPERATESBY ATTACHING TO THE FORWARD CABIN DOOR RESTRAINTWHEN THE DOOR IS IN THE FULLY OPEN POSITION.

CABIN DOOR HOLDOPEN DEVICE

(STOWED)

CLIP

Figure 4−10. Cabin Door Hold Open Device

CAUTION




4−16.ONE ENGINE INOPERATIVE TRAINING

TRAIN mode:

Placing an engine control switch in the TRAIN position will simulate a one engineinoperative (OEI) condition by resetting the selected engine’s governed speedto 92% NP, thereby putting the engine on standby while allowing single enginetraining on the opposite engine. In the event of an engine failure (or inadvertentswitching to IDLE) on the opposite engine, the engine in TRAIN will automatical-ly revert to 100% NP. Also, if the opposite engine control switch is placed inTRAIN both engines will revert to 100% NP.

NOTE: When an engine is placed into TRAIN the opposite engine will retain the 5 minuteTake−off Power engine parameter limiters and the IIDS does not rescale. Theresult is more realistic pilot OEI training, providing rotor droop in training if thepower requested is above the limiters as would happen in a real OEI condition.

IDLE mode:

If rescaling of the TORQUE and EGT displays and activation of the ENG OUTwarning is desired, the pilot should select IDLE instead of TRAIN for OEI train-ing. In the event the opposite engine should fail during this time the pilot mustselect FLY on the engine control switch to bring the good engine back to 100%NP.

NOTE: When operating with one engine in the IDLE mode OEI limits apply. OEI limitsare generally considered for ‘‘emergency use only’’ and excursions into thoselimits require recording in the engine log book and may increase themaintenance required. See Section VIII for recording and maintenance actionrequirements.

Precautions:

Pilots should consider such things as flight mode, gross weight, density altitudeand aircraft familiarity before conducting OEI training to avoid excursions intoOEI limits.

Recommended maximum takeoff weight for OEI training:

6000 LBS below 5000 Ft HD

5200 LBS at or above 5000 Ft HD

NOTE: For recommended Category A OEI training weights, refer to Section XI, Part IX.

4−17.FUEL SYSTEM

Capacities − Fuel System:

JET A: 1097 LBS; 498 kg; 161.3 U.S. gal; 611L total capacity

1078 LBS; 158.5 U.S. gal; 600L useable

JET B: 1048 LBS; 476 kg; 161.3 U.S. gal; 611L total capacity1030 LBS; 158.5 U.S. gal; 600L useable


Performance Data


S E C T I O N VPERFORMANCE DATA

TABLE OF CONTENTS

PARAGRAPH PAGE5−1. General 5−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−2. Noise Characteristics 5−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−3. Density Altitude Chart 5−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−1. Density Altitude Chart 5−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−4. Airspeed Calibration 5−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−2. Airspeed Calibration Curve 5−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−5. Best Rate of Climb Speed 5−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−3. Best Rate of Climb Speed (VY) 5−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−6. Rate of Climb and Descent − OEI 5−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−4. Single Engine Rate of Climb and Descent, From −36°C to 0°Cat VY, OEI MCP, and 3500 LB Gross Weight 5−9. . . . . . . . . . . . . . . . .


Figure 5−6. Single Engine Rate of Climb and Descent, From 0°C to −36°Cat VY, OEI MCP, and 4000 LB Gross Weight 5−11. . . . . . . . . . . . . . . . .






Figure 5−12. Single Engine Rate of Climb and Descent, From −36°C to −0°Cat VY, OEI MCP, and 5500 LB Gross Weight 5−17. . . . . . . . . . . . . . . . .


Figure 5−14. Single Engine Rate of Climb and Descent, at VY, OEI MCP, and 5750 LB Gross Weight 5−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Performance Data

FAA ApprovedReissue 1Original5−ii

PARAGRAPH PAGEFigure 5−15. Single Engine Rate of Climb and Descent, at VY,

OEI MCP, and 6000 LB Gross Weight 5−20. . . . . . . . . . . . . . . . . . . . . . .

Figure 5−16. Single Engine Rate of Climb and Descent, at VY, OEI MCP, and 6250 LB Gross Weight 5−21. . . . . . . . . . . . . . . . . . . . . . .

Figure 5−17. Single Engine Rate of Climb and Descent, at VY, OEI MCP,6500 LBS Gross Weight 5−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−7. Rate of Climb − AEO 5−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−18. Rate of Climb − AEO, at VY, MCP, 3500 Pounds Gross Weight 5−24.








Figure 5−26. Rate of Climb − AEO, at VY, TOP, 3500 Pounds Gross Weight 5−32. .








5−8. Hover Ceiling, AEO 5−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−34. Controllability Envelope and Azimuth Range for Crosswind Operations 5−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−35. Hover Ceiling, IGE, 3.5 Foot Skid Height, Standard Engine Inlet, Takeoff Power, Cabin Heat Off 5−42. . . . . . . .

Figure 5−36. Hover Ceiling, IGE, 3.5 Foot Skid Height, Standard Engine Inlet, Takeoff Power, Cabin Heat On 5−43. . . . . . . .

Figure 5−37. Hover Ceiling, OGE, Standard Engine Inlet, Takeoff Power, Cabin Heat Off 5−44. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−38. Hover Ceiling, OGE, Takeoff Power, Standard Engine Inlet, Cabin Heat On 5−45. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−39. Hover Ceiling, IGE, 3.5 Foot Skid Height, IPS Installed, Takeoff Power, Cabin Heat Off 5−46. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−40. Hover Ceiling, IGE, 3.5 Foot Skid Height, IPS Installed, Takeoff Power, Cabin Heat On 5−47. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Performance Data

FAA ApprovedReissue 1Original 5−iii/(5−iv blank)

PARAGRAPH PAGEFigure 5−41. Hover Ceiling, OGE, IPS Installed, Takeoff Power,

Cabin Heat Off 5−48. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−42. Hover Ceiling, OGE, IPS Installed, Takeoff Power, Cabin Heat On 5−49. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−9. Hover Ceiling, OEI 5−50. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−43. Hover Ceiling, OGE, Standard Inlet, 2.5 Minute OEI Power 5−51. . .

Figure 5−44. Hover Ceiling, OGE, IPS, 2.5 Minute OEI Power 5−52. . . . . . . . . . . . .

5−10. Height Velocity Diagram 5−53. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−45. Height Velocity Diagram 5−53. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−11. Power Assurance Check − Automatic 5−54. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−46. Power Assurance Check Menu 5−54. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

5−12. Power Assurance Check − Manual 5−56. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−47. Engine Torque Chart 5−58. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−48. EGT Chart 5−59. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 5−49. NG Chart − (NG read from Secondary IIDS Display) 5−60. . . . . . . . . .

Figure 5−50. NG Chart − (NG read from Third Level Power Check Menu) 5−61. . .


Performance Data




5−1

SECTION VPERFORMANCE DATA

5−1. GENERAL

This section contains baseline helicopter performance information as defined withincertain conditions such as airspeed, weight, altitude, temperature, wind velocityand engine power available. Data is applicable to the basic helicopter without anyoptional equipment installed unless otherwise noted.

5−2. NOISE CHARACTERISTICS

NOTE: No determination has been made by the Federal Aviation Administration that thenoise levels of this aircraft are or should be acceptable or unacceptable foroperation at, into, or out of, any airport.

The MD900 meets the FAR Part 36−J noise requirements at the certified maximumgross weight of 6500 LB for level flight at 0.9 VH.

MD900 NOISE CHARACTERISTICS − 6500 lb (2948 kg)

ENGINE: PW 207E

Configuration 0.9 VH(S.L. at 25°C)

118 KTAS

81.2 dBA

Clean aircraft, doors on,no external kits.


MD900 (902 Configuration with PW 207E)Performance Data


5−2

5−3. DENSITY ALTITUDE CHART

Description: The chart allows a quick estimation of the density altitude whenpressure altitude and OAT are known. This chart can also be used to determinetrue airspeed.

Use of Chart:

To determine density altitude, the pilot must know pressure altitude and outsideair temperature. Enter bottom of chart with known or estimated OAT, moveup to known pressure altitude line, move to left and note density altitude.

Pressure altitude is found by setting 29.92 (1013 mb) in Kolsman window± altimeter error.

To determine true airspeed convert indicated airspeed (IAS) to calibrated airspeed(CAS) utilizing the Airspeed Calibration Curve (Ref. Figure 5−2). Read valueon right of chart opposite known density altitude. Multiply CAS by this valueto determine true airspeed.

Examples:

Find density altitude for 6000 HP at −15°C:

Follow −15°C line to 6,000 ft pressure altitude line; read density altitude (3800ft).

Find density factor:

Read directly across from density altitude: (3800 ft). Note density factor of 1.058.

Find true airspeed:

130 KIAS = 127 KCAS (from Figure 5−2)127 KCAS � 1.058 = 134.4; round to 134 knots true airspeed.


Performance Data




5−3

−40 −30 −20 −10 0 10 20 30 40

−2000

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

DE

NS

ITY

ALT

ITU

DE

− F

EE

T

0.98

1.00

1.02

1.04

1.06

1.08

1.10

1.12

1.14

1.16

1.18

1.20

1.22

1.24

1.26

1.28

1.30

1.32

1.34

1.36

TEMPERATURE − °C

TEMPERATURE − °F

50 60

−40 −30 −20 −10 0 10 20 30 40 50 60 70 80 90 100 110F927−009

120 130 140

Figure 5−1. Density Altitude Chart




5−4

5−4. AIRSPEED CALIBRATION

Description: This charts show the difference between indicated and calibratedairspeeds.

Indicated airspeed (IAS) corrected for position error equals calibrated airspeed(CAS).

Use of chart: Use the chart as illustrated by the example. To determine calibratedairspeed, the pilot must know the indicated airspeed.

NOTE: The example below refers to Figure 5−2.

Example:

Wanted: Calibrated airspeed

Known: Indicated airspeed = 120 knots

Method: Enter the bottom of the chart at the indicated airspeed of 120 knots.Move up to the airspeed calibration line; move left and read 117 knots,calibrated airspeed.


Performance Data




5−5

F927−010

160

140

120

100

80

60

40

2016014012010080604020


CA

LIB

RA

TE

D A

IRS

PE

ED

− K

NO

TS

Figure 5−2. Airspeed Calibration Curve




5−6

5−5. BEST RATE OF CLIMB SPEED

Description: This chart shows the indicated airspeed to use for the best rate ofclimb at any given density altitude.

Use of Chart: Use the chart as illustrated by the example below.

Example:

Wanted: Best rate of climb

Known: Density altitude = 8,000 feet

Method: Enter the left side of chart at the known density altitude of 8,000 feet.Move to the right to the airspeed calibration curve and then directlydown to read 60 knots indicated airspeed (IAS) as the best rate of climbspeed.


Performance Data




5−7

F927−011

20 25 30 35 40 45 50 55 60 65 70 75 80


0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000D

EN

SIT

Y A

LTIT

UD

E −

FE

ET

Figure 5−3. Best Rate of Climb Speed (VY)




5−8

5−6. RATE OF CLIMB AND DESCENT − OEI

Description: These charts (Ref. Figure 5−4 thru Figure 5−17) show the rate ofclimb vs pressure altitude at maximum continuous OEI power at gross weightsranging from 3500 LB to 6500 LB at the best rate of climb speed.

NOTE: These charts based on an electrical load of 30%, heater off, and air-conditioningoff.

Use of Chart: The following example explains the correct use of the chart inFigure 5−4.

Use of Charts: Use the chart as illustrated by the example below.

Example:

Wanted: Rate of climb

Known: Pressure altitude = 4000 feet

Known: Outside air temperature = 0°C

Method: Enter the left side of chart (Ref. Figure 5−4) at the known pressurealtitude of 4000 feet. Move to the right to the 0°C temperature curveand then directly down to read rate of climb of approximately 1750feet per minute.


Performance Data




5−9

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

RATE OF CLIMB AT VY − FT/MINF927−012−11

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−200 0 200 400 600 800 1000 1200 1400 1600 1800

Continuous OEI Power, Vy, 3500 lb

0

−10

−20

−30

−36

OAT − �C

MAXIMUMOAT LIMIT

Figure 5−4. Single Engine Rate of Climb and Descent, From −36°C to 0°Cat VY, OEI MCP, and 3500 LB Gross Weight




5−10

F927−012−12A

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−200 0 200 400 600 800 1000 1200 1400 1600 1800 2000

40.5° OAT LIMIT

52° OAT LIMIT

30

40

50

20

10

0

−10


PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

RATE OF CLIMB AT VY − FT/MIN

OAT − �C



Performance Data




5−11

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

RATE OF CLIMB AT VY − FT/MINF927−012−2

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

0

−10

−20

−30

−36

OAT − �C

MAXIMUMOAT LIMIT


Figure 5−6. Single Engine Rate of Climb and Descent, From 0°C to −36°Cat VY, OEI MCP, and 4000 LB Gross Weight




5−12

F927−012−1A

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

40.5° OAT LIMIT

52° OAT LIMIT

30

40

50

20

10

0

−10

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

OAT − �C




Performance Data




5−13

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

0

−10

−36

OAT − �C


F927−012−3

Continuous OEI Power, Vy and 4500 lb

−20

−30

MAXIMUMOAT LIMIT





5−14

F927−012−4A

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600


PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

OAT − �C


52° OAT LIMIT

40.5° OAT LIMIT 30

40

50

20

10

0

−10



Performance Data




5−15

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

RATE OF CLIMB AT VY − FT/MIN F927−012−6

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

0

−10

−36

OAT − �C


MAXIMUMOAT LIMIT

−30

−20





5−16

F927−012−5A

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600


PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

OAT − �C


30

40

50

20

10

0

−10

40.5° OAT LIMIT

52° OAT LIMIT



Performance Data




5−17

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T


F927−012−8

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

0

−10

−36

OAT − �C

Continuous OEI Power, Vy and 5500 lb

−30

−20

Figure 5−12. Single Engine Rate of Climb and Descent, From −36°C to −0°Cat VY, OEI MCP, and 5500 LB Gross Weight




5−18

F927−012−7

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T


OAT − �C


30

40

50

20

10

0

−10

40.5° OAT LIMIT

52° OAT LIMIT



Performance Data




5−19

F927−012−13A

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T


OAT − �C


−30

−20

−36

30

40

50

20

10

0

−10

52° OAT LIMIT

40.5° OAT LIMIT

Figure 5−14. Single Engine Rate of Climb and Descent, at VY, OEI MCP,and 5750 LB Gross Weight




5−20

F927−012−9A

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T


OAT − �C


40.5° OAT LIMIT

52° OAT LIMIT

−36

−30

−20

30

40

50

20

10

0

−10

Figure 5−15. Single Engine Rate of Climb and Descent, at VY, OEI MCP, and 6000 LB Gross Weight


Performance Data




5−21

F927−012−14

−1000

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

−600 −400 −200 0 200 400 600 800 1000 1200 1400 1600

40.5° OAT LIMIT

52° OAT LIMIT

30

40

50

20

10

0

−10

−20

−36

−30

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T



OAT − �C

Figure 5−16. Single Engine Rate of Climb and Descent, at VY, OEI MCP, and 6250 LB Gross Weight




5−22

F927−107A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

−200 −100 0 100 200 300 400 500

−30°C

−36°C

−20°C

−10°C

0°C

10°C

20°C

30°C

5000 FT HD

40°C

41.5°COAT LIMIT

50°C

52°COAT LIMIT

THIS CHART BASED ON ELECTRICAL LOAD OF 30%,HEATER OFF, AND AIR−CONDITIONING OFF

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T

OAT



Figure 5−17. Single Engine Rate of Climb and Descent, at VY, OEI MCP,6500 LBS Gross Weight


Performance Data




5−23

5−7. RATE OF CLIMB − AEO

Description: These charts show the rate of climb vs pressure altitude at twin engine(AEO) MCP (Ref. Figure 5−18 thru Figure 5−24) or TOP (Figure 5−26 thruFigure 5−33) at the best rate of climb speed.

NOTE: These charts based on an electrical load of 30%, heater off, and air-conditioningoff.


Example:

Wanted: Rate of climb

Known: Pressure altitude = 3000 feet

Known: Outside air temperature = 20°C

Method: Enter the left side of chart at the known pressure altitude of 3000 feet.Move to the right to the 20°C temperature curve and then directly downto read rate of climb of 4200 feet per minute.




5−24

F927−013−7A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

1300 1500 1700 1900 2100 2300 2500 2700 2900 3100 3300 3500 3700 3900 4100 4300 4500

Rate of Climb − ft/min

Pre

ssu

re A

ltit

ud

e −

ft

MCP, VY, 3,500 lb

30

40

50

20

10

0

−10

−20

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C

Figure 5−18. Rate of Climb − AEO, at VY, MCP, 3500 Pounds Gross Weight


Performance Data




5−25

F927−013−1A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000 3200 3400 3600 3800


Pre

ssu

re A

ltit

ud

e −

ft

MCP, VY, 4000 lb

30

40

50

20

10

0

−10

−20

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C





5−26

F927−013−2A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800 3000 3200


Pre

ssu

re A

ltit

ud

e −

ft

30

40

50

20

10

0

−10

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C

MCP, VY, 4500 lb

−20



Performance Data




5−27

F927−013−3A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800


Pre

ssu

re A

ltit

ud

e −

ft

MCP, VY, 5000 lb

30

40

50

20

10

0

−10

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C

−20





5−28

F97−013−4A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400


Pre

ssu

re A

ltit

ud

e −

ft

OAT − �C

MCP, VY, 5500 lb

30

50

20

10

0

−10

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C

−20

40



Performance Data




5−29

F927−013−5A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200


Pre

ssu

re A

ltit

ud

e −

ft

MCP, VY, 6000 lb

30

50

20

10

0

−10

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C

−20

40





5−30

F927−013−6A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000


Pre

ssu

re A

ltit

ud

e −

ft

MCP, VY, 6250 lb

30

50

20

10

0

−10

−30

40.5° OAT LIMIT

52° OAT LIMIT

−36

OAT − �C

−20

40



Performance Data




5−31

F927−081B

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000


Pre

ssu

re A

ltit

ud

e −

ft

OAT −oC

52°OAT LIMIT

40.6° OAT LIMIT

14000 FT HDLIMIT

MCP, VY, 6500 lb

30

50

20

10

0

−10

−30

−36

−20

40





5−32

F927−013−8A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

1900 2100 2300 2500 2700 2900 3100 3300 3500 3700 3900 4100 4300 4500 4700 4900 5100


Pre

ssu

re A

ltit

ud

e −

ft

30

40

50

20

10

0

−10

OAT − �C

−20

−30/−36

40.5° OAT LIMIT

52° OAT LIMIT

TOP , VY, 3,500 lb

Figure 5−26. Rate of Climb − AEO, at VY, TOP, 3500 Pounds Gross Weight


Performance Data




5−33

F927−021−1A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

1300 1500 1700 1900 2100 2300 2500 2700 2900 3100 3300 3500 3700 3900 4100 4300 4500


Pre

ssu

re A

ltit

ud

e −

ft

TOP, VY, 4000 lb

50

40.5° OAT LIMIT

52° OAT LIMIT

30

40

20

10

0

−10

−20

−30/−36

OAT − �C





5−34

F927−021−2A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

900 1100 1300 1500 1700 1900 2100 2300 2500 2700 2900 3100 3300 3500 3700 3900


Pre

ssu

re A

ltit

ud

e −

ft

TOP, VY, 4500 lb

50

40.5° OAT LIMIT

52° OAT LIMIT

30

40

20

10

0

−10

−20

OAT − �C

−30/−36



Performance Data




5−35

F927−021−3A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

500 700 900 1100 1300 1500 1700 1900 2100 2300 2500 2700 2900 3100 3300


Pre

ssu

re A

ltit

ud

e −

ft

TOP, VY, 5000 lb

50

40.5° OAT LIMIT

52° OAT LIMIT

30

40

20

10

0

−10

−20

−30/−36

OAT − �C





5−36

F927−021−4A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600 2800


Pre

ssu

re A

ltit

ud

e −

ft

TOP, VY, 5500 lb

50

40.5° OAT LIMIT

52° OAT LIMIT

30

40

20

10

0

−10

−20

−30/−36

OAT − �C



Performance Data




5−37

F927−021−5A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600


Pre

ssu

re A

ltit

ud

e −

ft

TOP, VY, 6000 lb

50

40.5° OAT LIMIT

52° OAT LIMIT

30

40

20

10

0

−10

−20

−30/−36

OAT − �C





5−38

F927−021−6A

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

17000

18000

19000

20000

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200 2400


Pre

ssu

re A

ltit

ud

e −

ft

TOP, VY, 6250 lb

50

40.5° OAT LIMIT

52° OAT LIMIT

30

40

20

10

0

−10

−20

−30/−36

OAT − �C



Performance Data




5−39

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

13000

14000

15000

16000

300 500 700 900 1100 1300 1500 1700 1900 2100

Pre

ssu

re A

ltit

ud

e −

ft

OAT − oC

F927−021−7A

Rate of Climb and Descent − ft/min

TOP, VY, 6500 LB

50

30

40

20

10

0

−10

−20

−36

−30

52°OAT LIMIT

40.6° OAT LIMIT

14000 FT HDLIMIT





5−40

5−8. HOVER CEILING, AEO

Description:The hover ceiling charts (Ref. Figure 5−35 thru Figure 5−42) show the maximumhover weight capability, in ground effect (IGE) or out of ground effect (OGE), bothengines operating at take off power for known conditions of pressure altitude andoutside air temperature, or alternately, the maximum hover ceiling for a knowngross weight and outside air temperature.

Refer to Figure 5−34 for HIGE operations in crosswind conditions.

F927−146C

GROSS WEIGHT − LBS

120°

135°

270°

190°

80°

17 KTS

17 KTS

AZIMUTH RANGEB

B

0°

A

15 KTS

ÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔ

DE

NS

ITY

ALT

ITU

DE

− F

EE

T

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

4000 4200 4400 52005000 60006500

4600 4800 5400 5600 5800 6200 6400 66006250

12400

IGE HOVER OPERATION INWINDS OF 17 KNOTS HAVEBEEN DEMONSTRATED FOR ALLAZIMUTHS UP TO THIS LINE.

IGE HOVER OPERATION IN WINDSIN EXCESS OF 17 KNOTS HAVEBEEN DEMONSTRATED INAZIMUTH RANGE �C"(SEE FIGURE BELOW).

MAXIMUM SAFE WINDS FOR HOVER OPERATIONS DECREASE WITHINCREASING DENSITY ALTITUDE. TAKEOFF AND LANDING OPERATIONS INCALM WINDS OR HEADWINDS

C

C

IGE HOVER OPERATIONLIMITED TO 15 KNOTSWHEN WIND IS FROMAZIMUTH RANGE ‘A’, OR 17KNOTS WHEN WIND ISFROM AZIMUTH RANGE �B"(SEE FIGURE BELOW).

C

Figure 5−34. Controllability Envelope and Azimuth Range for Crosswind Operations


Performance Data




5−41

Separate hover ceiling charts are provided for helicopters equipped with either theinlet particle separator (IPS) or screen inlet and heater operation.The phrase, ‘‘A/C On’’ apply to MDHS P/N 900P7250302−101 air-conditioning instal-lation only.

NOTE: The charts are based on an electrical load of 15% per generator (30 amps pergenerator). Reduce/increase gross weight capability by 35 LB for each 10%increase/decrease in total load.

For many operations, a reduction in gross weight capability still allows the aircraftto operate at a maximum gross weight of 6500 LB. Follow the example shownbelow.

Use of Chart: The following example explains the correct use of the IGE Chartin Figure 5−35.

Example:

Wanted: Maximum gross weight for hover at 3.5 feet skid height at takeoff power.

Known: PA = 7000 feet; OAT = 30°C; cabin heat off and A/C on; 25% electricalload.

Method: Enter the chart at 30° OAT and move vertically to the 7000 PA curve(dashed lines). At this point, move directly to the left of the chart andread from the gross weight scale 6280 pounds.

Gross weight data above 6500 LB has been provided for calculationpurposes or external load operations only. Weights above 6500LB must be external and jettisonable.

The instructions for using the IGE hover ceiling charts also apply to the OGE hoverceiling charts.

CAUTION




5−42

F927−014−1C

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900

GR

OS

S W

EIG

HT

− L

B

13000

1400015000

12000

PRESSURE ALTITUDE − FEET

15300 HD

AIRCRAFT WITHOUTGENERATOR COOLING

MODIFICATION

16000

NOTE: GROSS WEIGHTS ABOVE 6500 LB PROVIDED FOR CALCULATION PURPOSES ONLY.

−50 −40 −30 −20 −10 0 10 20 30 40 50 60

11000

10000

9000

7000

6000

8000

5000

4000

3000

OAT °C

THIS CHART BASED ON WINDS 3KTS OR LESS AND 15% ELECTRICAL LOAD. FOR ELECTRICAL LOADS ABOVE/BELOW15%, DECREASE/INCREASE WEIGHT CAPABILITY 35 LBS PER 10% CHANGE IN ELECTRICAL LOAD.

REDUCE WEIGHT CAPABILITY 40 LBS WITH A/C ON

Figure 5−35. Hover Ceiling, IGE, 3.5 Foot Skid Height, Standard Engine Inlet, Takeoff Power, Cabin Heat Off


Performance Data




5−43

F927−014−2B

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900

GR

OS

S W

EIG

HT

− L

B

OAT °C−50 −40 −30 −20 −10 0 10 20 30 40 50 60

11000

10000

9000

7000

13000

1400015000

12000

16000

8000

MAXIMUM TEMPERATUREFOR CABIN HEAT ON


15300 HD


MODIFICATION

6000

5000


THIS CHART BASED ON WINDS 3KTS OR LESS AND 15% ELECTRICAL LOAD. FOR ELECTRICAL LOADSABOVE/BELOW 15%, DECREASE/INCREASE WEIGHT CAPABILITY 35 LBS PER 10% CHANGE IN ELECTRICAL LOAD.

Figure 5−36. Hover Ceiling, IGE, 3.5 Foot Skid Height, Standard Engine Inlet, Takeoff Power, Cabin Heat On




5−44

F927−015−1B

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900−40 −30 −20 −10 0 10 20 30 40 50 60

OAT °C

11000

10000

9000

7000

6000

13000

1400015000

12000

16000

8000

5000

4000

3000

2000

1000

0




NOTE: MAXIMUM INTERNAL GROSS WEIGHT 6500 LB. WEIGHTS IN EXCESS OF 6500 LB MUST BE EXTERNAL AND JETTISONABLE.

15300 HD

GR

OS

S W

EIG

HT

− L

B


MODIFICATION

Figure 5−37. Hover Ceiling, OGE, Standard Engine Inlet, Takeoff Power, Cabin Heat Off


Performance Data




5−45

F927−015−2B

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900−40 −30 −20 −10 0 10 20 30 40 50 60

11000

10000

9000

13000

1400015000

12000

16000

8000

OAT °C


7000



15300 HD

GR

OS

S W

EIG

HT

− L

B

6000

5000

4000

3000


MODIFICATION


Figure 5−38. Hover Ceiling, OGE, Takeoff Power, Standard Engine Inlet, Cabin Heat On




5−46

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900

F927−014−4B

THIS CHART BASED ON WINDS 3KTS OR LESS AND 15% ELECTRICAL LOAD. FOR ELECTRICAL LOADS ABOVE/BELOW 15%,DECREASE/INCREASE WEIGHT CAPABILITY 35 LBS PER 10% CHANGE IN ELECTRICAL LOAD.


GR

OS

S W

EIG

HT

− L

B

OAT °C−50 −40 −30 −20 −10 0 10 20 30 40 50 60


15300 HD

11000

10000

9000

7000

6000

13000

1400015000

12000

8000

5000

16000


MODIFICATION

2000

50003000

4000


Figure 5−39. Hover Ceiling, IGE, 3.5 Foot Skid Height, IPS Installed, Takeoff Power, Cabin Heat Off


Performance Data




5−47

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900

F927−014−5B

THIS CHART BASED ON WINDS 3KTS OR LESS AND 15% ELECTRICAL LOAD. FOR ELECTRICAL LOADS ABOVE/BELOW 15%,DECREASE/INCREASE WEIGHT CAPABILITY 35 LBS PER 10% CHANGE IN ELECTRICAL LOAD.

−50 −40 −30 −20 −10 0 10 20 30 40 50 60

GR

OS

S W

EIG

HT

− L

B

OAT °C


15300 HD

13000

1400015000

12000

16000


MODIFICATION11000

10000

9000

8000

7000


6000

5000


Figure 5−40. Hover Ceiling, IGE, 3.5 Foot Skid Height, IPS Installed, Takeoff Power, Cabin Heat On




5−48

F927−015−4B

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900−40 −30 −20 −10 0 10 20 30 40 50 60

OAT °C

GR

OS

S W

EIG

HT

− L

B

11000

10000

9000

7000

6000

13000

1400015000

12000

16000

8000

5000

4000

3000

2000

1000


0

15300 HD




MODIFICATION


Figure 5−41. Hover Ceiling, OGE, IPS Installed, Takeoff Power, Cabin Heat Off


Performance Data




5−49

F927−015−5B

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

6600

6700

6800

6900−50 −40 −30 −20 −10 0 10 20 30 40 50 60

11000

10000

9000

7000

6000

13000

140001500016000

8000

5000

4000

3000



15300 HD

12000

OAT °C

MAXIMUMTEMPERATUREFOR CABIN HEAT ON

GR

OS

S W

EIG

HT

− L

B


MODIFICATION

NOTE: MAXIMUM INTERNAL GROSS WEIGHT 6500 LB.. WEIGHTS IN EXCESS OF 6500 LB MUST BE EXTERNAL AND JETTISONABLE.

Figure 5−42. Hover Ceiling, OGE, IPS Installed, Takeoff Power, Cabin Heat On




5−50

5−9. HOVER CEILING, OEI

Description: These charts (Ref. Figure 5−43 and Figure 5−44) may be used to deter-mine hover performance in zero wind conditions for internal load operations orin headwind conditions during external load operations with one engine inoperative(emergency conditions) and the remaining engine at 2.5 minute power rating.

NOTE: Unless otherwise authorized by operating regulations, the pilot is not authorizedto credit more that 50 percent of the performance increase resulting from theactual favorable head wind increase.

NOTE: These charts are not to be used while conducting Category A takeoff and landingoperations.


Example 1: Zero wind

Wanted: Maximum gross weight for hover OGE at 2.5 minute OEI power.Known: HP = 4000 FT, OAT = 10°C

Method: Enter the chart at 10°C and move right to the 4000 HP curve. At thispoint move up and read from the gross weight scale, 5275 LB.

Example 2: Headwind

NOTE: It is essential that reliable wind information be available prior to determininghover. Additionally, only the lower limit of a gust spread may be used to determinehead wind credit.

Wanted: Maximum gross weight for hover OGE at 2.5 minute OEI power.Known: HP = 4000 FT, OAT = 10°C, 10 knot head wind

Method: Enter the chart at 10°C and move right to the 4000 HP curve. At thispoint move down to the 10 knot headwind line. From this point, move to theleft and read from the gross weight scale, 5475 LB.

Next, subtract 5275 LB (from example 1) from 5475 LB to determine the unfac-tored head wind performance increase of 200 LB. However, the pilot is authorizedto allow only 50 percent of the performance credit, resulting in a gross weightincrease to 5375 LB.


Performance Data




5−51

−1,000

5,000

8,0007,000

9,000

1,000

4,000

2,000

3,000

0

10,000

6,000

14,000

13,000

12,000

11,000

NOTE:

WIND SPEEDS AREUNFACTORED. APPLYFACTOR AS REQUIREDBY OPERATIONALRULES

F927−022−1A

THIS CHART IS BASED ON OEICONDITIONS, 2.5 MIN POWER

WIND FROM THE NOSE ±30DEGREES AND CABIN HEAT OFF

HEADWIND − KNOTS

GROSS WEIGHT

6700

3700

3900

4100

4300

4500

4700

4900

5100

5300

5500

5700

5900

6100

6300

6500

POUNDS

0−5

10

1520

25

30

GROSS WEIGHT − POUNDS

−50

−40

−30

−20

−10

0

10

20

30

40

50

60

OAT − °C


4000

4100

4200

4300

4400

4500

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500

3900

3800

3700

5600

5700


MODIFICATION

Figure 5−43. Hover Ceiling, OGE, Standard Inlet, 2.5 Minute OEI Power




5−52

−1,000

5,000

8,000

7,000

9,000

1,000

4,000

2,000

3,000

0

10,000

6,000

14,000

13,000

12,000

11,000

F927−022−2A

HEADWIND − KNOTSGROSS WEIGHT

3700

3900

4100

4300

4500

4700

4900

5100

5300

5500

5700

5900

6100

6300

6500

4000

4100

4200

4300

4400

4500

4600

4700

4800

4900

5000

5100

5200

5300

5400

5500


3900

3800

3700

−50

−40

−30

−20

−10

0

10

20

30

40

50

60

OAT − °C


POUNDS

0−510

152025

30

THIS CHART IS BASED ON OEICONDITIONS, 2.5 MIN POWER

WIND FROM THE NOSE ±30DEGREES AND CABIN HEAT OFF

NOTE:

WIND SPEEDS ARE UNFAC-TORED. APPLY FACTOR ASREQUIRED BY OPERATIONALRULES

5600

5700

6700


MODIFICATION

Figure 5−44. Hover Ceiling, OGE, IPS, 2.5 Minute OEI Power


Performance Data




5−53

5−10.HEIGHT VELOCITY DIAGRAM

ÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒÒ

ÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖÖ

F927−016

100

80

60

40

20

0

ALT

ITU

DE

− A

GL

(F

EE

T)


0 5 10 15 20 25 30

SMOOTH HARD SURFACE − WIND CALM

6251 TO 6500 LBAVOID AREA

CHART ‘‘A’’GROSS WEIGHT − LBS

DE

NS

ITY

ALT

ITU

DE

− F

EE

T

NOTE: IF THE COMBINATION OFGROSS WEIGHT ANDDENSITY ALTITUDE FALLIN THE SHADED REGIONOF CHART ‘‘B’’, THE‘‘AVOID AREAS’’ INCHART ‘‘A’’ APPLY.

CHART ‘‘B’’

140

120

6001 TO6250 LBAVOID AREA

ÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔÔ0

1000

2000

3000

4000

5000

6000

7000

5800 5900 6000 6100 6200 6300 6400 6500

Figure 5−45. Height Velocity DiagramThe clear area of Figure 5−45 Chart ‘‘B’’ represents density altitude/gross weightcombinations for which the height velocity diagram does not apply.

Safe landings and single engine fly−aways following an engine failure have beendemonstrated for the conditions stated below.

6000 LB gross weight at 7000 Ft HD and 6250 LB gross weight at 1400 Ft HD:

Safe landings following a vertical descent were demonstrated up to a 35 FTskid height.

Safe run−on landings were demonstrated up to a 90 FT skid height.

Fly−aways were demonstrated down to a 100 FT skid height.




5−54

5−11.POWER ASSURANCE CHECK − AUTOMATIC

The power assurance check provides a means for the pilot to determine, prior totake off, that each engine is capable of developing specification power.

NOTE: The primary purpose of this chart is its use as an engine performance trendingtool to aid in determining whether the engine is producing specification power,or if engine power deterioration has occurred. Power check data taken at regularintervals should be plotted to monitor trends in engine condition. Any trendindicating a reduction in engine performance should be investigated.

If desired, pilots can view the last power check under the IIDS POWER CHECKmenu or other previous power assurance checks in the TREND LOG underAIRCRAFT MONITOR menu.

NOTE: This power check procedure refers to the automated IIDS power check. If unableto perform the automated power check, use the manual power check methodfound in paragraph 5−12.

PERFORM POWER

ASSURANCE CHK

POWER CHECK

VIEW LAST POWER

ASSURANCE CHK

L PA CHK NG−X.X

L PA CHK EGT−XX.X

AUTOMATIC

RECORD DONE

L PA CHK NG−X.X

L PA CHK EGT−XX.X

R PA CHK NG−X.X

R PA CHK EGT−XX.X

R PA CHK NG−X.X

R PA CHK EGT−XX.X

NOTE: TO RETURN TO PREVIOUS HIGHER LEVEL − PRESS MENU

GND POWER CHK LFT ENG TQ XXX.X%

TIME 30 SEC

PRESS REC

RECORD DONE

RT ENG TQ XXX.X%

TIME 30 SEC

PRESS REC

AUTOMATIC

AUTOMATIC

AUTOMATIC

LT NG = XXX.X%

RT NG = XXX.X%

NOTE 1

NOTE 1: USED WHEN PERFORMING A MANUAL

POWER ASSURANCE CHECK.

NOTE2: PRESS REC KEY TO SAVE DATA IN TREND LOG;

MENU OR CLR KEY ABORTS FUNCTION

F927−017

TOP LEVEL SECOND LEVEL THIRD LEVEL FOURTH LEVEL

NOTE 2

Figure 5−46. Power Assurance Check Menu


Performance Data




5−55

HOW TO PERFORM THE CHECK:

NOTE: Power checks should be performed under the following conditions.1. Aircraft should be faced into the wind.2.Wind speed should not exceed 15 knots nor gust spread 5 exceed knots while

performing the check.3.Operate engine to be checked at 100% NP for five minutes to assure proper

operating temperatures are attained. 4. IPS and CABIN HEAT should be off and the generator load should be 10% or less.

� The engine to be checked should be at FLY.

� The other engine should be at IDLE or OFF.

� IPS and CABIN HEAT should be off and the generator load should be 10% or less.

� Select POWER CHECK top level menu on IIDS alphanumeric display.

� Press the ENT key 3 times to access the fourth level menu. LFT ENG TQ XXX.X%

TIME 30 SEC

Notice that the IIDS lists the left engine as the first engine to be checked. If the

the right engine is to be checked first, press the to access the

right engine menu. RT ENG TQ XXX.X%

TIME 30 SEC

� Stabilize engine torque at �3% of the ENG TQ value displayed for 30 seconds. TheIIDS provides a countdown from 30 seconds on the alphanumeric display duringdata acquisition. The countdown is started after the torque value is within the 3%range for more than 2 seconds.

NOTE: Counter will reset to 15 seconds if torque setting is not maintained within 3% for thelast 15 seconds of count down.

� After the IIDS calculates the performance margin of the selected engine, the

RECORD DONE

PRESS REC

menu is displayed and advises the pilot to press the REC

key to generate a trend log (Ref. Section VII) and to display the results of the powercheck on the alphanumeric display.

NOTE: If the power check fails, the IIDS displays a warning on the alphanumeric display.

� Lower collective and place engine control switch to IDLE

� After NP stabilizes, place other engine control switch to FLY.

� Press the to access the right engine menu; press the to

access the left engine menu.

� Repeat check for other engine.

NOTE: The engine torque value displayed should be approximately the same as the firstengine.




5−56

VIEWING THE PREVIOUS POWER CHECK:

� Select POWER CHECK top level menu on IIDS alphanumeric display.

� Press the ENT key once to access the second level menu. PERFORM POWER

ASSURANCE CHK

� Press the key to enter the next second level menu. VIEW LAST POWER

ASSURANCE CHK

� Press the ENT to view the last power check. L PA CHK NG−X.X

L PA CHK EGT−XX.X

� Press the key to view the results for the other engine. R PA CHK NG−X.X

R PA CHK EGT−XX.X

5−12.POWER ASSURANCE CHECK − MANUAL

HOW TO PERFORM THE CHECK:

� The engine to be checked should be at FLY.

� The other engine should be at IDLE or OFF.

� IPS and CABIN HEAT should be off and the generator load should be 10%or less.

� Record the IIDS OAT and pressure altitude.

� Use the Engine Torque Chart (Ref. Figure 5−47) to determine the torque valueto be utilized based on the OAT and pressure altitude recorded in the previousstep.

� Increase collective and stabilize at the predetermined torque value. After oneminute, record the EGT and NG from the IIDS.

� Use the EGT Chart (Ref. Figure 5−48) and the NG Chart (Ref. Figure 5−49)or Figure 5−50) to determine maximum values of EGT and NG for the specificconditions. Subtracting the recorded values from the maximum values willresult in the EGT and NG margins.

NOTE: The IIDS displays NG in tenths (ie. 91.6%) viewable at the third level of thePOWER CHECK menu.

� The power check is passed if both the EGT and NG margins are greater thanor equal to zero. If either the EGT or NG margin is negative, repeat the testallowing torque to stabilize for 5 minutes. If the EGT margin is still negative,then the power assurance check is failed. If only the NG margin is negativerefer to the Rotorcraft Maintenance Manual for additional testing and trouble-shooting.


Performance Data




5−57

EXAMPLE:

NOTE: This example assumes the inability to access the third level POWER CHECKmenu and therefor uses Figure 5−49 to determine maximum NG value.

Recorded from the IIDS: OAT = +30°CPressure Altitude = 2000 ft.

Utilizing the Engine Torque Chart (Ref. Figure 5−47) the power setting for the abovenoted conditions is determined to be:

Engine torque = 71%

Utilizing the EGT and NG Power Check Charts (Ref. Figure 5−48 and Figure 5−49)the maximum values for EGT and NG for the above noted conditions is determinedto be:

EGT = 791°C

NG = 92.5%

After stabilizing the torque at 71% for one minute you record the following EGTand NG readings from the IIDS:

EGT = 770°C

NG = 92%

Subtract the observed values of NG and EGT from the maximum values obtainedfrom the charts to determine the power check margins:

EGT = 791°C (from chart) minus 770°C (from IIDS) = 21°C (pass)

NG = 92.4% (from chart) minus 92% (from IIDS) = 0.5% (pass)




5−58

F927−018

−20 −10 0 10 20 30 40

35

40

45

50

55

60

65

70

75

80

EN

GIN

E T

OR

QU

E (

%)

AMBIENT TEMPERATURE (°C)

50

12000

6000

8000

14000

16000

10000

4000

2000

−30−36

SEA LEVEL

PRESSURE ALTITUDE (FEET)

Figure 5−47. Engine Torque Chart


Performance Data




5−59

500

550

600

650

700

750

800

850

900

−40 −20 0 20 40 60

PRESSURE ALTITUDE − FEET 4000

8000

SEA LEVEL

16000

12000

AMBIENT TEMPERATURE (°C) F927−019

Figure 5−48. EGT Chart




5−60

81

82

83

84

85

86

87

88

89

90

91

92

93

94

95

96

−40 −30 −20 −10 0 10 20 30 40 50

NG

− %

AMBIENT TEMPERATURE (‘C)


8000

16000

12000

4000

SEA LEVEL

F927−020−1

Figure 5−49. NG Chart − (NG read from Secondary IIDS Display)


Performance Data




5−61/(5−62 blank)

F927−020−2

83

84

85

86

87

88

89

90

91

92

93

94

95

96

97

−40 −30 −20 −10 0 10 20 30 40 50 60

AMBIENT TEMPERATURE (‘C)

NG

− %

PRESSUREALTITUDE − FEET

20000

16000

12000

8000

SEA LEVEL

4000

Figure 5−50. NG Chart − (NG read from Third Level Power Check Menu)

Weight andBalance Data


Original 6−iReissue 1

S E C T I O N V IWEIGHT AND

BALANCE DATATABLE OF CONTENTS

PARAGRAPH PAGE6−1. Weight and Balance Characteristics 6−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 6−1. Center of Gravity Limits 6−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−1. Center of Gravity Envelope. 6−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−2. Reference Coordinates 6−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−3. Station Diagram 6−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−4. Sample Weight and Balance Record 6−5. . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−5. Sample Weight and Balance Report 6−6. . . . . . . . . . . . . . . . . . . . . . . . . .

6−2. Load Limits and Balance Criteria 6−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−3. Equipment Removal or Installation 6−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 6−2. Cockpit, Cabin, and Baggage Compartment Doors Weight Data 6−7. .

Table 6−3. Cabin Doors Open Weight Data 6−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−4. Longitudinal Weight and Balance Determination:Passenger Configuration 6−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EXAMPLE I: Longitudinal CG Determination − Passenger 6−8. . . . . . . . . . . . . . . .

6−5. Longitudinal Loading of Cargo 6−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EXAMPLE II: Longitudinal CG Determination − Cargo 6−9. . . . . . . . . . . . . . . . . . .

6−6. Permissible Lateral Loadings − PassengerConfiguration 6−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EXAMPLE III: Lateral CG Determination − Passenger 6−10. . . . . . . . . . . . . . . . . . . .

6−7. Lateral Loading of Cargo 6−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

6−8. Internal Loading of Cargo 6−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EXAMPLE IV: Tiedown 500 pounds of cargo in the main cabin. 6−12. . . . . . . . . . . .

Table 6−4. Internal Cargo Loading 6−13. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−6. Cargo Restraint 6−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−7. Fuel Station Diagram − Jet−A 6−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 6−8. Fuel Station Diagram − Jet−B 6−16. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 6−5. Fuel Loading Table − Jet A (6.8 lb/U.S. gal) 6−17. . . . . . . . . . . . . . . . . . . .

Table 6−6. Fuel Loading Table − Jet B (6.5 lb/U.S. gal) 6−18. . . . . . . . . . . . . . . . . . . .

Table 6−7. Weight and Longitudinal Moments − Pilot, Passengers, Baggage 6−19.

Weight andBalance Data


Original6−iiReissue 1

PARAGRAPH PAGETable 6−8. Weight and Longitudinal Moments − Cargo 6−20. . . . . . . . . . . . . . . . . . . .

Table 6−9. Weight and Lateral Moments − Pilot and Passengers 6−21. . . . . . . . . . . .

Table 6−10. Weight and Lateral Moments − Cargo 6−22. . . . . . . . . . . . . . . . . . . . . . . .


MD900 (902 Configuration with PW 207E) Weight andBalance Data


SECTION VIWEIGHT AND BALANCE

DATA

6−1. WEIGHT AND BALANCE CHARACTERISTICS

The weight and balance characteristics are as follows:

Maximum weight on the landing gear with thruster extension: 6500 pounds.

Minimum Flying Weight: 3500 pounds.

Longitudinal Reference Datum: 199.3 inches forward of rotor hub centerline (rotorhub centerline is located at Station 199.3)

Cargo Deck Capacity: 1500 pounds not to exceed 115 pounds per square foot.

Baggage compartment limit (sta. 234.3 to 256.9): 500 pounds not to exceed 115lbs per square foot.

Ultimate load factors (cargo restraint): Forward: 17 G’sLateral: 8 G’s

Center of Gravity Limits:

NOTE: Lateral ‘‘+’’ is right of centerline ; lateral ‘‘−’’ is left of centerline when lookingforward.

Table 6−1. Center of Gravity Limits

Gross WeightLongitudinal C.G. Limit

(Sta−in.)Lateral C.G. Limit

(Sta−in.)

(lb) Forward Aft (+) Right, (−) Left

6500 196.0 202.3 +2.0; −2.0

6250 196.0 203.2 +2.0; −2.0

5100 196.0 206.0 +2.0; −2.0

*3500 196.0 206.0 +2.0; −2.0

Airspeed restrictions apply. Refer to Section II:

6250 196.0 202.1 +5.0; −2.0

5100 196.0 203.7 +5.7; −2.0

*3500 196.0 204.4 +6.0; −2.0

*Minimum flying weight.


MD900 (902 Configuration with PW 207E)Weight andBalance Data


F92−051C

LONGITUDINAL CG ENVELOPE

LATERAL CG ENVELOPE

WHEN OPERATING IN THEEXPANDED CG REGION OF CHARTA, THE MAXIMUM LONGITUDINALC.G. LIMIT, AS DEPICTED BY THEDASHED LINE IN CHART B, APPLIES.

3000

3500

4000

4500

5000

5500

6000

6500

−3 −2 −1 0 1 2 3 4 5 6 7

3000

3500

4000

4500

5000

5500

6000

6500

194 196 198 200 202 204 206 208

GR

OS

S W

IGH

T −

LB

SG

RO

SS

WIG

HT

− L

BS

CHART A: LATERAL C.G. STATION (IN.)

EXPANDEDCG LIMITS

CHART B: LONGITUDINAL C.G. STATION (IN)

5100 LBS

NORMAL CG LIMITS

NORMAL CG LIMITS

Figure 6−1. Center of Gravity Envelope.




+15.85

−15.85

STA 130.7 STA 173.0

0.0

+19.0

−19.0

STA 213.0 STA 245.6

CG OF PILOT ORCOPILOT/PASSENGER

CG REAR FACINGPASSENGERS

CG FWD FACINGPASSENGERS

CL OF BAGGAGE

COMPARTMENT

F92−052

CG CABIN

STA 193.0

0.0

Figure 6−2. Reference Coordinates




F92−053A

60 40 20 0 −20 −40 −60

220

200

180

160

140

120

100

80

60

WL 106FLOOR

BL 24BEAM

BL 8.5BEAM

−5050

220

200

180

160

140

120

100

80

60

STA 155.5FRAME

STA 230.5FRAME

3° 16"

STA 292.817WL 147

WL 159ROOF DECK

60 80 100 120 140 180 200 220 240 260 280 300 320 340 360 380 400 420 440 460

STA 199.25WL 207.97

160

5° 0"

WL 106FLOOR

JACKINGPOINTS

480

Figure 6−3. Station Diagram




BA

SIC

WE

IGH

T A

ND

BA

LA

NC

E R

EC

OR

D(C

ON

TIN

UO

US

HIS

TO

RY

OF

CH

AN

GE

S IN

ST

RU

CT

UR

E O

R E

QU

IPM

EN

T A

FF

EC

TIN

G W

EIG

HT

AN

D B

ALA

NC

E)

AIR

CR

AF

T M

OD

EL

MD

900

SE

RIA

L N

UM

BE

R90

0−00

0XX

X

MD

Hel

icop

ters

, In

c.

DAT

E

ITE

M N

O.

INO

UT

RE

GIS

TR

AT

ION

NU

MB

ER

N9X

XX

XP

AG

E 4

OF

4

DE

SC

RIP

TIO

N O

F A

RT

ICLE

WE

IGH

T C

HA

NG

ER

UN

NIN

G T

OTA

LB

AS

IC A

IRC

RA

FT

AD

DE

D (

+)R

EM

OV

ED

(−)

WE

IGH

TLO

NG

AR

MW

EIG

HT

LON

GA

RM

For

m H

OQ

014

(rev

5/0

0)

12/2

3/xx

01/0

9/xx

01/0

9/xx

XA

CT

UA

L B

AS

IC W

EIG

HT

RE

VIS

ED

CA

LCU

LAT

ED

BA

SIC

WE

IGH

T

FIX

ED

BA

LLA

ST

IN N

OS

E5.

087

.643

8

3277

.821

0.2

6891

03

210.

432

72.8

6886

65

F92−054A

OR

MO

CIF

ICA

TIO

NLA

TA

RM

LON

GM

OM

EN

TLA

TM

OM

EN

TW

EIG

HT

LON

GA

RM

LAT

AR

MLO

NG

MO

ME

NT

LAT

MO

ME

NT

LAT

AR

MLO

NG

MO

ME

NT

LAT

MO

ME

NT

0.4

1465

0.4

1309

Figure 6−4. Sample Weight and Balance Record




Weighed by

Model Serial No. Reg. No.

J. Doe

MD900 900−000XX N92XXX 12/23/XXDate

AIRCRAFT ACTUAL WEIGHT

FUELOIL, ENGINE LHOIL, ENGINE RHOIL, TRANSMISSIONHYDRAULIC FLUID

X

EMPTY FULL

X

X

X

X

F92−187A

220

200

180

160

140

120

100

80

60

STA 155.5FRAME

STA 230.5FRAME

3° 16"

STA 292.817WL 147

WL 159ROOF DECK

60 80 100 120 140 180 200 220 240 260 280 300 320 340 360 380 400 420 440 460

STA 199.25WL 207.97

160

5° 0"

WL 106FLOOR

JACKINGPOINTS

480

NOTE: IN A LEVEL ATTITUDE, MAIN ROTOR MAST IS TILTED 3 DEG. FORWARD.

WEIGHING POINTAVE. SCALE

READINGLBS

TARE ORCALIB. CORR.

LBS

NETWEIGHT

LBS

LONGITUDINALARM

IN

LATERALARM

IN

LONGITUDINALMOMENTIN−LBS

LATERALMOMENTIN−LBS

Forward 869.7 0.0 869.7 154.0 −9.3 133929 −8066

Aft Right 1289.5 0.0 1289.5 233.0 23.3 300454 29981

Aft Left 887.5 0.0 887.5 233.0 −23.3 206788 −20634

TOTAL (AS WEIGHED) 3046.7 210.4 0.4 641170 1280

Less : Surplus Weight (See Table 1) −1.8 189.4 −10.4 −346 19

Plus: Missing Required Equipment (See Table 1) 222.6 198.0 0.0 44071 0

TOTAL − BASIC WEIGHT 3267.4 209.6 0.4 684895 1299

Figure 6−5. Sample Weight and Balance Report




6−2. LOAD LIMITS AND BALANCE CRITERIA

The load limits and balance conditions are as noted in Table 6−1.

Do not exceed these limitations at any time during flight.

Use the helicopter Basic Weight as recorded in the Basic Weight and Balance Recordinserted in this section to perform all weight and balance computations. Basic Weightincludes oil, hydraulic fluid, and unusable fuel.

6−3. EQUIPMENT REMOVAL OR INSTALLATION

Removal or addition of equipment must be entered on the repair and alterationreport form, FAA 337, in accordance with Federal Air Regulations which shall thenbecome part of the helicopter log book file.

Record the weight and balance effects of these changes in the Basic Weight andBalance Record inserted in this section.

Use the balance and station diagrams shown as an aid for weight and balancechanges.

Use the following tables to assist in determining weight and balance effects withdoors opened or removed.

Table 6−2. Cockpit, Cabin, and Baggage Compartment Doors Weight Data

ITEMWEIGHT

(LB)LONGITUDINAL

STATION(ARM)

LATERALSTATION

(ARM)

MOMENT(IN−LB)

Longitudinal Lateral

Cockpit doors (2) 24.0 132.9 �30.6 3190 �367

Cabin doors (2) 32.2 196.0 �31.2 6311 �502

Baggage door (1) 9.4 269.0 0 2529 0

Table 6−3. Cabin Doors Open Weight Data

ITEM

WEIGHT(LB)

LONGITUDINALSTATION

(ARM)

LATERALSTATION

(ARM)

MOMENT(IN−LB)


Cabin doors (2) 32.2 248.0 �31.2 7986 �502

Note: At minimum flying weight (3500 LBS) the CG shifts 0.48 inch aft with cabin doorsopen.

At maximum gross weight (6250 LBS) the CG shifts 0.27 inch aft with cabin doors open.




6−4. LONGITUDINAL WEIGHT AND BALANCE DETERMINATION:PASSENGER CONFIGURATION

To determine that the gross weight and longitudinal center of gravity (fore andaft) for a given flight are within limits, proceed as follows.

Obtain aircraft basic weight and moment from the Weight and Balance Record in-serted in this section.

Determine weights and moments of useful load items (Ref. Figure 6−2).

Add above items.

Determine corresponding center of gravity for gross weight by dividing total momentby gross weight. This computation must be done with zero fuel and with missionfuel gross weight (Ref. EXAMPLE I: ).

NOTE: If loadings are not symmetrical about the aircraft centerline, determine lateralCG’s as described in Paragraphs 6−6 and 6−7.

EXAMPLE I: Longitudinal CG Determination − Passenger

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)

Basic Weight (from Figure 6−4) 3272.8 688665

Pilot 170 130.70 22219

Copilot/Passenger 170 130.70 22219

Passenger − Rear Facing R/H 170 173.0 29410

Passenger − Rear Facing L/H 170 173.0 29410

Passenger − FWD Facing R/H 170 213.0 36210

Passenger − FWD Facing L/H 170 213.0 36210

1. Zero Fuel WeightAdd: Fuel (Jet−A)

4292.8994.0 191.1

864343189953

2. Gross Weight 5286.8 1054296

Calculation of Longitudinal CG

CG at Zero Fuel Weight:

Moment at Zero Fuel Weight=

864343= 201.3

Zero Fuel Weight 4292.8

CG at Gross Weight:

Moment at Gross Weight=

1054296= 199.4

Gross Weight 5286.8

NOTE: The CG’s fall within the limits specified in Table 6−1; therefore, the loading meetsthe longitudinal CG limits.




6−5. LONGITUDINAL LOADING OF CARGO

The large aft compartment of the Model 900 provides great flexibility in the varietyof cargo loads it can accommodate.

To determine the gross weight and center of gravity for a given flight are withinlimits, proceed as follows.Obtain the Basic Weight and Moment from the Weight and Balance Record (Ref.Figure 6−4).

Establish the weight of cargo load.

Determine the location of the cargo longitudinal CG (Ref. Figure 6−3)

Obtain the cargo moment:

Cargo Moment = Cargo Weight X Cargo CGPerform weight and balance as previously described for passenger configura-tion.

EXAMPLE II: Longitudinal CG Determination − Cargo

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)


Pilot 170 130.7 22219

Copilot/Passenger 170 130.7 22219

Cargo 750 190.0 142500


4362.8300.0 187.0

87560356100

2. Gross Weight 4662.8 931703

Calculation of Longitudinal CG



875603= 200.7


CG at Gross Weight:


931703= 199.8

Gross Weight 4662.8

NOTE: The CG’s fall within the limits specified in Table 6−1; therefore, the loading meetsthe longitudinal CG limits.




6−6. PERMISSIBLE LATERAL LOADINGS − PASSENGERCONFIGURATION

Safe operation of this helicopter requires that it be flown within established lateralas well as longitudinal center of gravity limits.

It is therefore imperative that lateral center of gravity control be exercised.

All combinations of internal loadings are permissible if gross weight, longitudinal,and lateral center of gravity considerations permit.

To determine the gross weight and center of gravity for a given flight are withinlimits, proceed as follows.Obtain the basic weight and longitudinal moment from The Basic Weight and Bal-ance Record (Ref. Figure 6−4).

For pilot and passenger longitudinal and lateral center of gravity stations, see

Figure 6−2.

EXAMPLE III: Lateral CG Determination − Passenger

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)


Pilot 170 +15.85 2695

Passenger − Rear Facing R/H 170 +19.00 3230

Passenger − FWD Facing R/H 170 +19.00 3230


3782.8500.0

−−10464

0

2. Gross Weight 4282.8 10464

Calculation of Lateral CG



10464= 2.77


CG at Gross Weight:


10464= 2.44

Gross Weight 4282.8

NOTE: The CG’s fall outside the limits specified in Table 6−1; therefore, the loadingdoes not meet the lateral CG limits.




6−7. LATERAL LOADING OF CARGO

To determine the gross weight and lateral center of gravity for a given flight arewith limits, proceed as follows.

Find weight of load.

Determine lateral location (station) of load center of gravity.

Measure load distance from aircraft (centerline) lateral station zero), right(+) : left (−).

Obtain the lateral load moment as follows.

Lateral moment = weight X lateral station (or use Table 6−10).

Perform weight and balance as previously described for longitudinal CG determina-tions.

6−8. INTERNAL LOADING OF CARGO

The following instructions should be followed when carrying internal cargo.

Restrain the cargo from shifting by using the correct number of tiedowns in accor-dance with Table 6−4.

Locate restraint loops in accordance with Figure 6−6.

NOTE: Cargo carried in the baggage compartment shall not be higher than 36 inches.

To assure that cargo is properly secured, refer to Table 6−4.

The numbered tiedown location is located in the far left column of Table 6−4with their respective restraint values in the six columns to the right.

Locate the cargo tiedown numbers for all of the tiedowns that you will be usingin the respective cargo areas (main cabin or baggage compartment).

Add the restraint values for each of the tiedowns in each of the three directions(forward, left and right).

If the sum of restraint values in each of the three directions equals or exceeds theweight of the cargo, then the cargo is sufficiently restrained.

NOTE: 1. Cargo should be centered in the cabin or baggage compartment.2. Do not load cargo outside the perimeter defined by the cargo tiedown fittings.




EXAMPLE IV: Tiedown 500 pounds of cargo in the main cabin.

LATERAL

TIEDOWN No. FORWARD LEFT RIGHT

1 −−− 220 −−−

14 −−− −−− 220

2 20 40 −−−

13 20 −−− 40

4 130 40 −−−

11 130 −−− 40

5 120 220 −−−

10 120 −−− 220

____ ____ ____

TOTAL 540 520 520

Since all three values exceed the weight of the cargo (500 pounds), the cargo issufficiently restrained.




Table 6−4. Internal Cargo Loading

TIE−DOWN LOCATION RESTRAINT VALUE/POUNDS OF CARGO

TIE−DOWNNUMBER

FUSELAGESTATION

LATERALSTATION

MAIN CABINRESTRAINT DIRECTION

BAGGAGE COMPARTMENTRESTRAINT DIRECTION

FORWARDLATERALLEFT (−)

LATERALRIGHT (+)

FORWARDLATERALLEFT (−)

LATERALRIGHT (+)

1 156.8 −27.0 220

2 174.9 −25.0 20 40

3 193.0 −25.0 130 20

4 211.1 −25.0 130 40

5 229.2 −27.0 120 220

6 229.2 −11.0 50 240

7 229.2 −8.0 50 240

8 229.2 8.0 50 240

9 229.2 11.0 50 240

10 229.2 27.0 120 220

11 211.1 25.0 130 40

12 193.0 25.0 130 20

13 174.9 25.0 20 40

14 156.8 27.0 220

15 156.8 11.0 240

16 156.8 8.0 240

17 156.8 −8.0 240

18 156.8 −11.0 240

19 232.9 −21.6 90 185

20 251.0 −24.8 120 135 110

21 233.3 0.0 85 85

22 257.9 0.0 20 120 120

23 232.9 21.6 90 185

24 251.0 24.8 120 135 110

25 230.5 −24.6 110

26 230.5 24.6 110

27 269.0 −17.5 105

28 269.0 17.5 105




CARGO RESTRAINT LOCATION

1

2

3

4

5 6 7 8 9 10

12

13

1415161718

19

20

21 23

24

25 26

27 28

FWD

LEFT RIGHT

11

F92−056

TIEDOWNS 25 AND 26 ARE ‘‘D’’RINGS LOCATED AT WL 154.5

TIEDOWNS 19 THRU 28 ARE‘‘D’’ RINGS. TIE DOWNS 27 AND28 ARE LOCATED AT WL 155.0

22

Figure 6−6. Cargo Restraint




183 184 185 186 187 188 189 190 191FUSELAGE STATION CG

0

25

50

75

100

125

150

175

200

225

250

275

300

325

350

375

400

425

450

475

500

525

550

575

600

625

650

675

700

725

750

775

800

825

850

875

900

925

950

975

1000

FU

EL

WE

IGH

T −

PO

UN

DS

F92−057−1

NOTES:

WEIGHTS AND MOMENTS BASED ON JET−A FUEL

2. TOTAL WEIGHT OF FUEL IS DEPENDENT UPONTHE SPECIFIC GRAVITY AND TEMPERATUREVARIATION SHOULD BE ANTICIPATED IN GAUGE READINGS WHEN TANKS ARE FULL.

3. FUEL CG VARIES WITH QUANTITY

(ASTM D−1655) AT 6.8 POUNDS PER U.S. GALLON

1025

1050

1075

1100

Figure 6−7. Fuel Station Diagram − Jet−A




F92−057−2

FU

EL

WE

IGH

T −

PO

UN

DS

183 184 185 186 187 188 189 190 191

FUSELAGE STATION CG

0

25

50

75

100

125

150

175

200

225

250

275

300

325

350

375

400

425

450

475

500

525

550

575

600

625

650

675

700

725

750

775

800

825

850

875

900

925

950

975

1000

NOTES:WEIGHTS AND MOMENTS BASED ON JET B FUEL

2. TOTAL WEIGHT OF FUEL IS DEPENDENT UPONTHE SPECIFIC GRAVITY AND TEMPERATUREVARIATION SHOULD BE ANTICIPATED IN GAUGE READINGS WHEN TANKS ARE FULL.

3. FUEL CG VARIES WITH QUANTITY

(ASTM D−1655) AT 6.5 POUNDS PER U.S. GALLON

1025

1050

Figure 6−8. Fuel Station Diagram − Jet−B




Table 6−5. Fuel Loading Table − Jet A (6.8 lb/U.S. gal)

VOLUME U.S. GALLONS

WEIGHTPOUNDS

STATIONINCHES

MOMENTIN−LBS

10 68 184.0 1251115 102 184.4 1881320 136 184.9 2514425 170 185.3 3150330 204 185.7 37888

35 238 186.1 4429740 272 186.5 5072845 306 186.9 5718150 340 187.2 6365355 374 187.5 70142

60 408 187.9 7664765 442 188.2 8316670 476 188.4 8969775 510 188.7 9623780 544 188.9 102787

85 578 189.2 10934390 612 189.4 11590595 646 189.6 122470

100 680 189.8 129038105 714 189.9 135607

110 748 190.1 142176115 782 190.2 148744120 816 190.3 155309125 850 190.4 161872130 884 190.5 168431

135 918 190.6 174986140 952 190.7 181537145 986 190.8 188033150 1020 190.8 194626155 1054 190.9 201165160 1088 190.9 207710

NOTES:1. TOTAL WEIGHT OF FUEL IS DEPENDANT UPON THE SPECIFIC

GRAVITY AND TEMPERATURE. VARIATION SHOULD BE ANTICI−PATED IN GAUGE READINGS WHEN TANKS ARE FULL.

2. FUEL CG VARIES WITH QUANTITY.3. MAXIMUM USEABLE FUEL QUANTITY IS 994 LBS.4. MAXIMUM USEABLE FUEL QUANTITY IS 1078 LBS. WITH RANGE

EXTENDER




Table 6−6. Fuel Loading Table − Jet B (6.5 lb/U.S. gal)

VOLUME U.S. GALLONS

WEIGHTPOUNDS

STATIONINCHES

MOMENTIN−LBS

10 65 184.0 1195915 98 184.4 1798320 130 184.9 2403525 163 185.3 3011330 195 185.7 36216

35 228 186.1 4234240 260 186.5 4849045 293 186.9 5465850 325 187.2 6084555 358 187.5 67048

60 390 187.9 7326665 423 188.2 7949770 455 188.4 8573975 488 188.7 9199280 520 188.9 98252

85 553 189.2 10451990 585 189.4 11079195 618 189.6 117067

100 650 189.8 123345105 683 189.9 129624

110 715 190.1 135903115 748 190.2 142181120 780 190.3 148457125 813 190.4 154730130 845 190.5 161000

135 878 190.6 167266140 910 190.7 173528145 943 190.8 179786150 975 190.8 186040155 1008 190.9 192290160 1040 190.9 198538

NOTES:1. TOTAL WEIGHT OF FUEL IS DEPENDANT UPON THE SPECIFIC

GRAVITY AND TEMPERATURE. VARIATION SHOULD BE ANTICIPATEDIN GAUGE READINGS WHEN TANKS ARE FULL.

2. FUEL CG VARIES WITH QUANTITY.3. MAXIMUM USEABLE FUEL QUANTITY IS 950 LBS. 4. MAXIMUM USEABLE FUEL QUANTITY IS 1030 LBS. WITH RANGE

EXTENDER




Table 6−7. Weight and Longitudinal Moments − Pilot, Passengers, Baggage

PASSENGERWEIGHT

(LBS)

PILOT OR COPILOT/PASSENGER STA 130.7

REAR FACING PASSENGER

STA 173.0

FWD FACINGPASSENGER

STA 213.0

MOMENT(IN−LB)

MOMENT(IN−LB)

MOMENT(IN−LB)

100 13070 17300 21300

120 15684 20760 25560

140 18298 24220 29820

160 20912 27680 34080

180 23526 31140 38340

200 26140 34600 42600

220 28754 38060 46860

240 31368 41520 51120

BAGGAGE(LBS)

AFTBAGGAGESTA 245.6 BAGGAGE

(LBS)

AFTBAGGAGESTA 245.6

MOMENT(IN−LB)

MOMENT(IN−LB)

100 24560 320 78592

120 29472 340 83504

140 34384 360 88416

160 39296 380 93328

180 44208 400 98240

200 49120 420 103152

220 54032 440 108064

240 58944 460 112976

260 63856 480 117888

280 68768 500 122800

300 73680




Table 6−8. Weight and Longitudinal Moments − Cargo

WEIGHT(LBS)

MOMENT (IN−LB)

STATION 160 STATION 180 STATION 200 STATION 220 STATION 240

100 16000 18000 20000 22000 24000

120 19200 21600 24000 26400 28800

140 22400 25200 28000 30800 33600

160 25600 28800 32000 35200 38400

180 28800 32400 36000 39600 43200

200 32000 36000 40000 44000 48000

220 35200 39600 44000 48400 52800

240 38400 43200 48000 52800 57600

260 41600 46800 52000 57200 62400

280 44800 50400 56000 61600 67200

300 48000 54000 60000 66000 72000

320 51200 57600 64000 70400 76800

340 54400 61200 68000 74800 81600

360 57600 64800 72000 79200 86400

380 60800 68400 76000 83600 91200

400 64000 72000 80000 88000 96000

420 67200 75600 84000 92400 100800

440 70400 79200 88000 96800 105600

460 73600 82800 92000 101200 110400

480 76800 86400 96000 105600 115200

500 80000 90000 100000 110000 120000




Table 6−9. Weight and Lateral Moments − Pilot and Passengers

PASSENGERWEIGHT

(LBS)

PILOT OR COPILOT/PASSENGER

STA. ±15.85*

REAR FACINGPASSENGERSTA. ±19.00*

FWD FACINGPASSENGERSTA. ±19.00*

MOMENT(IN−LB)

MOMENT(IN−LB)

MOMENT(IN−LB)

100 1585 1900 1900

110 1744 2090 2090

120 1902 2280 2280

130 2061 2470 2470

140 2219 2660 2660

150 2378 2850 2850

160 2536 3040 3040

170 2695 3230 3230

180 2853 3420 3420

190 3012 3610 3610

200 3170 3800 3800

210 3329 3990 3990

220 3487 4810 4810

230 3646 4370 4370

240 3804 4560 4560

250 3963 4750 4750

260 4121 4940 4940

270 4280 5130 5130

280 4438 5320 5320

290 4597 5510 5510

300 4755 5700 5700

*Indicated moments are + (right lateral) and − (left lateral).




Table 6−10. Weight and Lateral Moments − Cargo

WEIGHT(LBS)

MOMENT (IN−LB)

LATERAL STATION ±5 in.*




20 100 200 300 400

40 200 400 600 800

60 300 600 900 1200

80 400 800 1200 1600

90 450 900 1350 1800

100 500 1000 1500 2000

110 550 1100 1650 2200

120 600 1200 1800 2400

130 650 1300 1950 2600

140 700 1400 2100 2800

150 750 1500 2250 3000

160 800 1600 2400 3200

170 850 1700 2550 3400

180 900 1800 2700 3600

l90 950 1900 2850 3800

200 1000 2000 3000 4000

210 1050 2100 3150 4200

220 1100 2200 3300 4400

230 1150 2300 3450 4600

240 1200 2400 3600 4800

250 1250 2500 3750 5000

260 1300 2600 3900 5200

270 1350 2700 4050 5400

280 1400 2800 4200 5600

290 1450 2900 4350 5800

300 1500 3000 4500 6000

*Indicated moments are + (right lateral) and − (left lateral).

Systems Description



S E C T I O N V I ISYSTEMS DESCRIPTION

TABLE OF CONTENTS

PARAGRAPH PAGE7−1. Helicopter Exterior Description 7−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−1. Helicopter − Major Components 7−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−2. Fuselage 7−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−2. Door Opening Decals − Exterior 7−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−3. Tailboom and Empennage 7−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−4. Landing Gear 7−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−3. Landing Gear 7−7. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−5. Main Rotor System 7−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−4. Main Rotor System 7−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−6. Flight Controls 7−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−5. Cyclic Controls Subsystem 7−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−6. Collective Controls Subsystem 7−13. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−7. Upper Flight Controls Subsystem 7−14. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−8. Anti−Torque Controls Subsystem (Sheet 1 of 3) 7−16. . . . . . . . . . . . . . . .



Figure 7−9. VSCS Control Subsystem 7−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−7. Hydraulic Systems 7−20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−10. Hydraulic System Installation 7−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−11. Hydraulic System Block Diagram 7−22. . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−12. Rotor Brake System 7−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−8. Propulsion System 7−24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−13. PW207E Engine Installation 7−24. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−14. Powerplant − Components 7−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−15. Drive System (Sheet 1 of 2) 7−26. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−15. Drive System (Sheet 2 of 2) 7−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−9. Engine Air Intake and Inlet Particle Separator (IPS) 7−28. . . . . . . . . . . . . . . . . . . . . . .

Figure 7−16. Engine Air Intake 7−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−10. Engine Power Management System 7−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Systems Description


Original7−iiReissue 1

PARAGRAPH PAGE7−11. Fuel System 7−31. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−17. Fuel System Schematic 7−32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−18. IIDS Fuel System Display 7−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−12. Fire Extinguishing System 7−34. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−19. Fire Extinguishing System 7−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−13. Electrical System 7−36. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−20. Battery Power and External Power Subsystem Block Diagram 7−36.

Figure 7−21. Battery Power, External Power, and DC Power Component Locator 7−37. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−14. Environmental Control 7−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−22. Heat/Defog System Schematic 7−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Figure 7−23. IIDS System Monitoring 7−42. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−24. IIDS Display Brightness Control 7−45. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−25. Alphanumeric Display 7−47. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 7−1. Automatic Alphanumeric Display Warning/Caution/Advisory Messages 7−47. . . . . . . . . . . . . . . . . . . . . . . .

7−16. IIDS Data Storage 7−49. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7−17. Balance Monitoring System 7−52. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−26. Balance Monitoring System Installation 7−52. . . . . . . . . . . . . . . . . . . . .

7−18. IIDS Menu Structures 7−53. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−27. IIDS Top Level Menus 7−53. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−28. Time Summary 7−54. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−29. Balance Monitor, Main Rotor Balance 7−55. . . . . . . . . . . . . . . . . . . . . . .

Figure 7−30. Balance Monitor, Run M/R Measurements 7−56. . . . . . . . . . . . . . . . . . .

Figure 7−31. Balance Monitor, Main Rotor Configuration 7−57. . . . . . . . . . . . . . . . . .

Figure 7−32. Balance Monitor, Main Rotor Solution Options 7−58. . . . . . . . . . . . . . .

Figure 7−33. Balance Monitor, Display M/R Solution 7−59. . . . . . . . . . . . . . . . . . . . . .

Figure 7−34. Balance Monitor, M/R Track 7−60. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−35. Balance Monitor, NOTARR Balance 7−61. . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−36. Balance Monitor, NOTARR Data 7−62. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−37. Balance Monitor, Spectrum 7−63. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−38. Balance Monitor, BMS Fault Log 7−64. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−39. Balance Monitor, BMS Version Log 7−65. . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−40. Balance Monitor, BMS Maintenance 7−66. . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−41. Aircraft Monitor, Exceedance Log Menu 7−67. . . . . . . . . . . . . . . . . . . . .

Systems Description


Original 7−iii/(7−iv blank)Reissue 1

PARAGRAPH PAGEFigure 7−42. Aircraft Monitor − Trend Log 7−68. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−43. Aircraft Monitor, Fault Log Menu 7−69. . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−44. Aircraft Monitor − IIDS Setup 7−70. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−45. Fuel Calibration 7−71. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−46. Set Engine Parameters 7−72. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 7−47. Set Time/Date 7−72. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


MD900 (902 Configuration with PW 207E) Systems Description


SECTION VIISYSTEMS DESCRIPTION

7−1. HELICOPTER EXTERIOR DESCRIPTION

Design features:

� Category A performance capabilities

� Cockpit with outstanding field of view

� All composite fuselage with expanded aluminium foil embedded in skin for light-ning protection

� Wide (52 in/132 cm), sliding cabin doors for loading bulky cargo

� Crash resistant fuel cell

� Built−in steps and work platforms for maintenance

� NOTAR� anti−torque system

� H−type empennage with twin vertical stabilizers

� Five−bladed main rotor with swept blade tips

� Hingeless low drag main rotor hub

� Optional engine inlet air particle separator

� On−board systems monitoring and computerized track and balance

The patented NOTAR� anti−torque system provides many benefits. It results inlow noise by locating the fan in the fuselage and eliminating the conventional noisytail rotor, provides outstanding safety because there is no exposed tail rotor, andimproved directional controllability over that of the conventional tail rotor helicopter.

The five−bladed main rotor is designed for outstanding performance and flying quali-ties. Vibration in the passenger spaces is minimized by the incorporation of thefive blades and the unique dynamically−tuned ‘‘static mount" that supports therotor and transmission. The swept tips on rotor blades improve performance andreduce main rotor noise. Interior noise is minimized by using an acoustic noiseattenuating support for the transmission gearbox, and acoustic insulation in theceiling and sidewalls of the cabin.


MD900 (902 Configuration with PW 207E)Systems Description


MAIN ROTOR BLADEASSEMBLY

ENGINERIGHTHAND

ENGINELEFT HAND

EMPENNAGEASSEMBLY

UPPER COWLINGAND FAIRINGS

TRANSMISSIONASSEMBLY TAILBOOM

ASSEMBLY

ANTI−TORQUETHRUSTER

BAGGAGECOMPARTMENTDOOR

ANTI−TORQUEASSEMBLY

CABIN DOOR

LANDING GEARASSEMBLY

COCKPITDOOR

FLIGHTCONTROLS

FUSELAGE STRUCTUREASSEMBLY

COCKPITDOOR

CABIN DOOR

F92−058

Figure 7−1. Helicopter − Major Components




The composite flexbeam main rotor hub replaces the normal hinges with a fiberglass/epoxy flexbeam that twists and bends to accommodate the blade motions. It, andthe elastomeric lead/lag dampers, are located within the elliptical pitchcase for alow drag hub that is composed of a minimum number of parts.

The empennage includes a fixed horizontal stabilizer and two controlled verticalstabilizers that provide directional stability.

The screened NOTAR� inlet is on the top of the cowling, between the engines andaft of the rotor. In this location it is protected from dust and debris, and is shapedto direct NOTAR� fan noise up and away from observers on the ground, thus helpingto minimize noise.

The cabin floor is approximately three feet above the ground. This provides spaceunder the fuselage for the energy absorbing landing gear to deflect, and room inthe lower fuselage for the 161.3 gallon fuel cell. A convenience step is providedon the right side of the fuselage for entering and departing.

Step/handholds and fold−out work platforms are built into the sides of the fuselage,forward and aft of the cabin doors, for easy access to equipment located on theengine and transmission decks.

Two tiedown fittings are positioned high on the sides of the fuselage in line withthe forward edge of the cabin doors, and one on the fuselage centerline just abovethe baggage compartment door. Fabric socks are used to capture the blade tipsfor tying them to the landing gear crosstubes.

7−2. FUSELAGE

The fuselage contains the cockpit; cabin; baggage compartment; fuel cell; NOTAR�

fan, support, and ducts; and avionics equipment. The rotor/transmission support,engines, and systems equipment are mounted on the top, and the landing gearon the bottom. The fuselage structure has an aluminum upper deck, main frames,and anti−plowbeams under the cockpit, with graphite/epoxy skins, keel beams,cockpit framing, floors, and doors.

The fuselage is one of three components that contribute to an integrated systemsapproach to the MD Explorer’s hard landing energy absorbing concept. The othersare the landing gear and crew/passenger seats. This approach has served well inthe OH−6A, AH−64A, and MD500 helicopters.

The cabin has an open flat floor from the front of the copilot’s station throughthe cabin and to the back of the baggage compartment area. With the seats removed,the entire floor area is usable for loading cargo.

Space is provided in the nose for the battery; under the cockpit floor and in thebaggage compartment for avionics equipment; and under the baggage compartmentfloor for air conditioning equipment. The single fuel cell is mounted in the bellyof the fuselage surrounded by bulkheads fore and aft, and keel beams to the sides.




Entry Doors:

Hinged cockpit doors, sliding cabin doors, and a hinged baggage compartment doorprovide access. The cockpit doors have door release handles that allow the doorsto be removed (Ref. Section VIII). The windows in the cabin doors are easily remov-able and the meet Transport Category emergency exit size criteria.

The cockpit door handles have four positions and main cabin door handles havethree positions:

F927−024CABIN DOOR OPENING DECALCOCKPIT DOOR OPENING DECAL

KEYLOCK

SAFELOCK

OPEN

SLAM

SAFELOCK

OPEN

LOCK

SLIDING

DOOR

Figure 7−2. Door Opening Decals − ExteriorThe rotor/transmission mount consists of an eight−legged metal truss that sup-ports the mast base and the static mast. The transmission gearbox mounts beneaththe mast base and the rotor turns on the static mast tube on a set of tapered rollerthrust bearings. Two of the truss tubes on the right side of the aircraft are removablefor transmission maintenance.

Graphite/epoxy cowlings and access doors on top of the fuselage enclose the equip-ment located there.

Saddle mounts in the lower fuselage clamp the forward and aft landing gear cross-tubes in place.

Lightning protection for the graphite/epoxy skins is provided by expanded alumi-num foil molded into the surface, with all components electrically bonded together.Electromagnetic pulse protection (EMP) is provided by the aluminum structure,the expanded aluminum foil on the graphite skins, and the shielding of individualelectric/avionics systems components and wiring.




7−3. TAILBOOM AND EMPENNAGE

Anti−torque, directional control, and yaw stability is provided by the NOTAR�

fan, circulation control tailboom, the thruster, and the horizontal and verticalstabilizers with VSCS.

The NOTAR� fan is driven directly from the main transmission. The fan is locatedin the aft fuselage, and supplies pressurized air to the tailboom (pressure ratio= 1.02 to 1.12). Its blade pitch and the thruster nozzle rotational positions are oper-ated by the anti−torque pedals.

The circulation control tailboom is a hollow graphite/epoxy cylinder that boltsto the aft end of the fuselage and supports the horizontal and vertical stabilizers,tail bumper, and the thruster. The tailboom directs the pressurized air to the thrusterwhile allowing some air to flow out of the two slots along its right side. This arrange-ment creates a significant side force on the tailboom as a result of the circulationflow around the tailboom while it is immersed in the main rotor downwash. Theremainder of the side force required for directional control is produced by airflowout of the controllable direct jet thruster at the end of the tailboom.

The empennage consists of the horizontal stabilizer with upper and lower moveablevertical stabilizers located at each tip. The horizontal and vertical surfaces are graph-ite/epoxy. The horizontal stabilizer has an inverted NACA 2412 airfoil with a fixedincidence of −1 degree. A trailing edge Gurney tab is installed above and belowthe airfoil to balance aerodynamic moments. The vertical stabilizers have a hybridNACA 23012/NACA 0012 airfoil cambered toward the right side of the helicopter.

The vertical stabilizers are controlled in incidence by electro−mechanical actuatorslocated within the horizontal stabilizer that operate in response to collective pitchinputs. Both vertical stabilizers also respond to the Vertical Stabilizer Control Sys-tem (VSCS) to function as a yaw damper.

To minimize tail vibration, the horizontal stabilizer attaches to the top of the tailboomwith an energy absorbing mount that is hinged along a fore−and−aft axis at theright side, and connected by an elastomeric damper on the left side.

Lightning protection is provided by a strip of aluminum foil bonded onto thesurface of the tailboom, expanded aluminum foil co−cured onto the empennage sur-faces, and jumpers to form a continuous electrical path to the fuselage.




7−4. LANDING GEAR

The landing gear (Ref. Figure 7−3) supports the helicopter when it is in contactwith the ground. The landing gear can withstand loads encountered during landing,ground handling, and provides a stable platform to prevent ground resonance.

The landing gear primarily absorbs normal landing forces, with the capabilitiesto absorb severe landing forces during overload conditions. The landing gear dimen-sions are based on the required minimum roll−over and minimum pitch−over angles.A minimum angle of 27 degrees is maintained from the center of gravity (CG) locationto the skid−to−ground contact point. The landing gear consists of the following compo-nents:

Forward and Aft Crosstubes − Provide energy absorbing capabilities during nor-mal or severe landings.

Forward and Aft Saddle Assemblies − Provide a means to attach the crosstubeassemblies to the fuselage attachment points.

Side Stop Clamp Assemblies − Prevent side movement of the crosstube assemblies.

Forward Spacer Fittings − Forward attachments for the skid tubes and forwardcrosstube assembly.

Skid Tubes − Provide landing gear−to−ground contact points.

Damper Assemblies − Aft attachments for the skid tubes and aft crosstube assem-bly.

Each damper has a reservoir fluid level indicator that is a rotating shaft whichshows through a 120� pie shaped window. When the reservoir is filled, the windowshows green with a very thin wedge of red showing to the first notch on thehousing. The thin wedge of red shows the reservoir is not completely full, toallow for fluid expansion.

Ensure fluid level in reservoir is within limits.

Reservoir is near empty, when the window shows red and should be serviced(RMM. Section 12−00−00).




LANDING GEAR DAMPER

RESERVIOR FLUIDLEVEL INDICATOR

FIRST NOTCH

SECOND NOTCH

GREENRED

EMPTY FULL

F92−060

RESERVOIR INDICATOR CLOCKING TYPICAL

SKID TUBE

FORWARDSPACER FITTING

AFTSADDLE ASSEMBLY

AFTCROSSTUBE

SIDE STOPCLAMP ASSEMBLY

PLUG

DAMPERASSEMBLY

AFTABRASION STRIP

FORWARDCROSSTUBE

FORWARDSADDLE ASSEMBLY

STEP

FORWARDABRASION STRIP

GROUND HANDLINGATTACH POINTS

MID ABRASIONSTRIP

Figure 7−3. Landing Gear




7−5. MAIN ROTOR SYSTEM

The main rotor is a five−bladed, fully articulated hingeless flexbeam system. Therotor diameter is 33.83 ft (10.34 m) with a blade chord of 10 in (25 cm). At its nominal100 percent rotational speed (NR), the rotor runs at 392 rpm (695 feet/second tipspeed).

The flexbeam is primarily a unidirectional fiberglass/epoxy, y−shaped member thatconnects the blade to the rotor hub, and twists and bends to accommodate the blademotions, resisting centrifugal force while transmitting drive torque to the blade.The five flexbeams attach to the hub by five bolts.

The pitchcase is a hollow, elliptically shaped graphite/epoxy tube that surroundsthe flexbeam and is attached to both the flexbeam and the blade at its outboardend by a pair of expandable−bushing bolts. The pitchcase provides flapwise, chord-wise, and torsional stiffness to the inboard end of the blade and serves to transmitthe feathering control motions to the blade. The pitchcase is attached to the hubat its inboard end by the elastomeric snubber/damper that provides centering forflapping and feathering motions, and by a combination spring/damper restraintfor chordwise motion to eliminate ground resonance. An elastomeric bumper isbonded to the flexbeam halfway along its length to bear against the inside of thepitchcase and restrict a flexbeam bending oscillation that would otherwise occurduring spin−up and shut−down of the rotor.




OUTBOARDABRASION STRIP

TRIM TABASSEMBLY

MAINROTOR

HUB

PITCHCASE

FLEXBEAM

ROTOR BLADE

INBOARDABRASION STRIP

ROTOR BLADERETENTION BOLTS

CENTERINGBEARING

PITCH CHANGEHORN

DRIVEPLATE

DAMPER

SCISSORS

LOWERHUB

FLEXBEAMBUMPER

UPPERHUB

DRIVERING

F92−061

Figure 7−4. Main Rotor SystemThe hub consists of two machined aluminum plates with a steel spacer betweenthem. The plates are grooved to accept the flexbeams and are bolted together withthe same bolts that attach the flexbeams. The hub mounts to the static mast bya pair of grease lubricated, tapered roller bearings. A splined drive plate bolts tothe top of the hub and is driven by the main rotor shaft that rotates inside themast.

This static mast rotor support configuration has been used successfully in theOH−6A, AH−64A, and MD500 helicopters and is incorporated into the MD Explorerfor three reasons:

Vibration Control − fuselage/mast/rotor structure is tuned dynamically for mini-mum vibration.

Reduced transmission weight − gearcase is not required to support rotor loads.




Safety − if the drive shaft should break, the rotor remains mounted to the mastby its two bearings for a safe autorotation landing.

The main rotor blade is made from fiberglass/epoxy with a hollow leading edgespar and a Nomex honeycomb−filled trailing edge. It has a theoretical twist of −10degrees; and the high performance airfoil tapers in thickness from 12 percent atit’s inboard end to 9.5 percent at the tip. The outboard 14 in. (36 cm) of the bladeplanform has a parabolic swept back taper. A 8 in. (20 cm) long by 3/4 in. (20 mm)chord trim tab is centered on the 77 percent radius station. Two pockets in thebottom of the blade near the tip are provided for installing blade balance weights.

A titanium abrasion strip protects the inboard, constant−chord portion of the bladewhile an electroformed nickel abrasion strip is fitted outboard. A polyurethane sheetprotects the under side of the blade outboard.

The MD Explorer has a built−in track and balance system for the main rotor andfor the NOTAR� fan blades that operates through the Integrated Instrument DisplaySystem (IIDS).

Lightning protection is afforded by a continuous electrical path from blade tipto rotor mast, and so on into the fuselage. This consists of the metal abrasion stripon the blade, expanded aluminum foil co−cured onto the surface of the pitchcase,dual jumpers across all joints, and twin carbon brushes for hub−to−mast continuity.

7−6. FLIGHT CONTROLS

The flight controls provide a means of controlling blade pitch of the main rotorin flight and during ground operations. The helicopter is equipped with dual pilotcontrols.

The flight controls integrate pilot inputs from the cyclic, collective, and anti−torquesubsystems. The cyclic and collective control stick inputs are mechanically linkedto the upper flight controls for longitudinal, lateral, and vertical control. The anti−torque pedal inputs are transmitted to the NOTAR� fan and direct jet thrusterfor directional control. The flight controls consist of the cyclic controls, collectivecontrols, upper flight controls, anti−torque controls, and vertical stabilizer controlsubsystems.

The cyclic controls subsystem controls helicopter pitch and roll attitudes (longitu-dinal and lateral control). The cyclic controls move the upper flight controls to cycleincreases or decreases in the rotor blades angle of attack in a cyclic manner aroundthe rotor azimuth. The result is a change in the helicopter pitch and/or roll attitude.

The cyclic control subsystem consists of the following components:

Cyclic stick assembly − Provides pilot control of helicopter pitch and/or roll atti-tude. This cyclic stick mount places the stick grip at its highest point abovethe floor when it is farthest aft − it moves down as it moves forward. This allowsthe pilot to rest his/her forearm on his/her thigh throughout all flight modesfor very comfortable flying.




Longitudinal and lateral trim actuators − Allow the pilot to position the cyclicas required during flight and while on the ground.

Longitudinal linkages − Allow for cyclic input to the main rotor blades for helicop-ter pitch control.

Longitudinal servoactuator − Hydraulically transfers longitudinal linkage inputsto position the upper flight controls.

Lateral linkages − Allow for cyclic input to the main rotor blades for helicopterroll control.

Lateral servoactuator − Hydraulically transfers lateral linkage inputs to positionthe upper flight controls.

The collective controls subsystem controls helicopter lift (vertical control) andthrust. As the collective stick assembly is moved, control linkages increase or de-crease the rotor blades angle of attack.

The collective pitch system includes two automatic control features:

Conventional ‘‘anticipatory" circuit into the Engine Electronic Controls (EEC) toprepare them for an upcoming change of power demanded by the changing collectivepitch position, and vertical stabilizer incidence angle change (VSCS).

The collective control subsystem consists of:

Collective stick assembly − Provides pilot control of helicopter lift.

Collective friction unit − Allows collective stick assembly resistance to varyfrom 5−25 lb (2.27−11.34 kg).

Collective friction release switch − Allows the pilot to release collective stickassembly resistance.

Collective linkages − Allows the pilot to transmit collective input to the upperflight controls.

Collective servoactuator − Hydraulically transfers collective linkage inputs tothe upper flight controls.




LONGITUDINALBELLCRANKASSEMBLY

CYCLIC CONTROLSTICK BOOT

CYCLIC BASEASSEMBLY

CYCLIC STICKASSEMBLY

LATERALBRACKET

ASSEMBLY

LATERALBELLCRANKASSEMBLY

AFT COCKPITLONGITUDINALCONTROL ROD

ASSEMBLY

LATERALCONTROL

RODASSEMBLY

LONGITUDINAL CLOSET-CONTROL ROD ASSEMBLY

LATERAL GRADIENTSPRING ASSEMBLY

LONGITUDINAL CONTROLSCONTROL ROD

ASSEMBLY

LATERALTRIM ACTUATOR

LATERAL CONTROLSBRACKET ASSEMBLY

LONGITUDINAL TRIM ACTUATORCRANK ASSEMBLY

LONGITUDINAL GRADIENTSPRING ASSEMBLY

EXPANDABLEDIAMETER

BOLT ASSEMBLY

REFUPPER DECK

DUAL LATERAL ROD ENDBALL BEARING

COCKPIT LATERAL CONTROLSTUBE ASSEMBLY

DUALLONGITUDINALCONTROL ROD

ASSEMBLY

LONGITUDINALBRACKET ASSEMBLY

LATERALCONTROLSCONTROL RODASSEMBLY

LATERALBELLCRANKASSEMBLYLONGITUDINAL

TRIM ACTUATOR ASSEMBLY

LATERAL TRIM ACTUATORCRANK ASSEMBLY

F92−062−1

Figure 7−5. Cyclic Controls Subsystem




PILOTCOLLECTIVE STICK

ASSEMBLY

COLLECTIVECONTROL STICK BOOT

COLLECTIVEFRICTION UNIT

COLLECTIVEBRACKET ASSEMBLY

COLLECTIVE STICKBRACKET ASSEMBLY

DETENT MODULEASSEMBLY

DETENT MODULEMOUNTING BRACKET

COLLECTIVECONTROL ROD ASSEMBLY

CONTROL BRACKETASSEMBLY

COLLECTIVEBRACKET ASSEMBLY

COLLECTIVE CONTROLROD ASSEMBLY

COLLECTIVEBELLCRANK ASSEMBLY

LONGITUDINAL/COLLECTIVEFOD COVER

UPPER DECKCOLLECTIVE CONTROLROD ASSEMBLY

COLLECTIVE HYDRAULICSERVOACTUATOR

COLLECTIVE FRICTIONRELEASE SWITCH

COLLECTIVE INTERCONNECTCONTROL ROD ASSEMBLY

INTERCONNECTCABLE ASSEMBLY


COPILOTCOLLECTIVE STICK ASSEMBLY

SENSORBRACKET

ASSEMBLY

SENSORLINK ASSEMBLY

SENSOR COLLECTIVEPOSITION BELLCRANKASSEMBLY

POTENTIOMETERCLAMP

POTENTIOMETERS


F92−062−2

Figure 7−6. Collective Controls Subsystem




Main Rotor Controls:

The main rotor mechanical control system uses conventional pushrods and bell-cranks under the cockpit floor; in the forward, right hand cockpit/cabin bulkhead;and in the cabin ceiling to transmit the control motions to the dual tandem hy-draulic actuators that operate the rotor control mixer and the swashplate.

SWASHPLATEASSEMBLY

MIXERASSEMBLY

SCISSORS DRIVELINK ASSEMBLY

COLLECTIVE DRIVELINK ASSEMBLY

LATERAL ANTI−TORQUEDRIVE LINK ASSEMBLY

ROTOR CONTROLPITCH LINK ASSEMBLY

F92−062−3

Figure 7−7. Upper Flight Controls Subsystem




Anti−torque Controls:

The anti−torque pedals are adjustable fore and aft and include an adjustablefriction device. They operate through a pushrod/bellcrank system and a singlehydraulic actuator to control the rotation of the direct jet thruster and changethe blade pitch angle of the NOTAR� fan to maintain constant air pressurein the tail boom as the thruster nozzle opens and closes. The hydraulic actuatoroperates the NOTAR� fan blade pitch through a pushrod/bellcrank/cam linkage,and the thruster rotation through a push/pull type cable along the length ofthe tailboom and a local tension cable loop at the thruster. The pedals do notcontrol the vertical stabilizers.

Attached to the lower directional crank assembly is the pedal anticipator. Thepedal anticipator provides the EEC’s an indication of impending anti−torquefan pitch change, which allows the EEC’s to anticipate an increase in powerdemand. The pedal anticipator also allows the IIDS to display and record pedalposition.




F92−062−4

DIRECTIONALINTERCONNECT

CONTROL ROD ASSEMBLY

PILOT DUAL CONTROLDIRECTIONAL

PEDAL ASSEMBLY

COPILOT DUAL CONTROLDIRECTIONAL PEDAL ASSEMBLY

RIGHTHEEL REST

SUPPORTLEFTHEEL RESTSUPPORT

HEEL RESTASSEMBLY

DIRECTIONALCONTROL ROD ASSEMBLY

LOWER CLOSETDIRECTIONALBELLCRANK ASSEMBLY

AFT DIRECTIONALCONTROL ROD ASSEMBLY

HEEL RESTASSEMBLY

REF UPPER DECKDIRECTIONAL

BELLCRANK ASSEMBLY

PEDALADJUSTMENT

HANDLE

PEDALCRANK

ASSEMBLY

DIRECTIONAL PEDALLINK ASSEMBLY

PEDALANTICIPATOR

Figure 7−8. Anti−Torque Controls Subsystem (Sheet 1 of 3)




F92−062−5A

NOTAR® FAN LINKAGE

INNER BELLCRANKASSEMBLY

DIRECTIONALBRACKET

ASSEMBLY DIRECTIONAL CABLEATTACH BRACKET

DIRECTIONALBELLCRANKASSEMBLY

DIVERTERPLATE ASSEMBLY

TO THRUSTERCONTROL

DIRECTIONAL CONTROLSCONTROL ROD ASSEMBLY

DIRECTIONALCONTROL CABLE

ASSEMBLY

DIRECTIONAL CONTROLSCONTROL ROD

ASSEMBLY

DIRECTIONALCONTROL ROD

ASSEMBLY

SPLITTERASSEMBLY

DIRECTIONALBRACKET ASSEMBLY


FODCOVER

CONTROLBRACKET ASSEMBLY


ANTI−TORQUESERVO ACTUATOR

DIRECTIONALBRACKET ASSEMBLY

UPPERDECK


ASSEMBLY


ASSEMBLY

OUTER BELLCRANKASSEMBLY

NOTAR® FAN INPUT FORCELIMITING CONTROL ROD

ASSEMBLY

NOTAR® FAN INPUT FORCELIMITING CONTROL ROD





F92−062−6

AFTTHRUSTER CONTROL

CABLE ASSEMBLY

THRUSTERSTATIONARYCONE ASSEMBLY

THRUSTER CONTROLDRUM ASSEMBLY

THRUSTER DRUMBRACKET ASSEMBLY

THRUSTER CONTROLSECTOR ASSEMBLY

THRUSTER CONTROLROD ASSEMBLY

THRUSTER BUILDUPASSEMBLY

ROTATING CONEASSEMBLY

TAILBOOMASSEMBLY

DIRECTIONAL CONTROLCABLE ASSEMBLY

THRUSTER CONTROLROD ASSEMBLY

VIEW ROTATED





The Vertical Stabilizer Control System (VSCS) operates the incidence of thevertical stabilizers through two electro−mechanical actuators, one for the left stabi-lizer and one for the right stabilizer. One portion of the system is a fly−by−wireactuator of stabilizer incidence as a function of collective pitch stick position. It’spurpose is to provide an anticipation that a power change is occurring to preventrotor droop and to maximize the anti−torque contribution of the stabilizers at highspeed thereby minimizing power required by the fan − leaving more power availablefor the main rotor. The second portion of the system is a fly−by−wire yaw dampingfunction that uses yaw gyro/lateral accelerometer signals to impose a supplementaryincidence on both vertical stabilizers. Instrumentation/control includes a dual indica-tor on the instrument panel to show incidence angle of the two vertical stabilizers;a LEFT STAB FAIL, RIGHT STAB FAIL, or TOTAL STAB FAIL yellow CAUTIONannunciator on the IIDS alphanumeric display; and two OFF/ON/TEST ‘‘L VSCSR’’ switches on the utility panel, and a ‘‘YAW SYNC’’ switch located on the collectivecontrol module (Ref. Section IV). The ‘‘YAW SYNC’’ switch allows the pilot to resetthe VSCS to operate around the current lateral acceleration and yaw rate. Thisfeature is useful when transitioning from hovering to forward flight, and when transi-tioning from a turn to level flight or from level flight into a turn.

VERTICAL STABL R

VERTICAL STABL R

VSCS INDICATOR

F92−063

LEFT VSCSCONTROL UNIT

YAW RATEGYRO

COLLECTIVECONTROL POSITION

TRANSDUCERS

YAWRATEGYRO

LEFTVERTICALSTABILIZER

RIGHTVERTICAL

STABILIZER

RIGHT LINEARACTUATOR

LEFTLINEAR

ACTUATOR

RIGHT VERTICALSTABILIZER LINKAGES

LEFT VERTICALSTABILIZER LINKAGES

RIGHT LATERALACCELEROMETER

RIGHT VSCSCONTROL UNIT

Figure 7−9. VSCS Control Subsystem




7−7. HYDRAULIC SYSTEMS

Flight Controls:

The helicopter is equipped with two hydraulic systems for operation of the flightcontrols. Under certain conditions, the main rotor control loads are such thatthey require at least one hydraulic system operating at all times; hence, thedual system for safety. However, the aircraft can be flown in a minimally degradedcondition with the anti−torque actuator depressurized.

The system is powered by two variable displacement hydraulic pumps mountedon and driven by the main transmission, has a reservoir/manifold for each systemplaced on opposite sides of the upper fuselage deck, and has three tandem actua-tors, one for each cyclic pitch function and one for collective pitch of the mainrotor.

The #1 system operates only the main rotor controls while the #2 system operatesthe main rotor controls and also the NOTAR�

anti−torque control system.

The main rotor actuators are mounted forward of the main rotor while the anti−torque actuator is mounted in the cabin ceiling just aft of the right hand cabindoor.

A hand pump option is installed for use in servicing the hydraulic systems inthe field.

The two systems normally operate at 500 psi each for a total system pressureof 1000 psi. If pressure in one system should drop to less than 400 psi, the othersystem automatically compensates by increasing its pressure to maintain a totalsystem pressure of 1000 psi nominal. A yellow caution annunciator,‘‘1 HYD’’ or ‘‘HYD 2’’, illuminates on the IIDS caution/warning display and acaution message is displayed on the alphanumeric display when the affectedsystem’s pressure falls below 250 psi.




PRESSURETRANSDUCER

SYSTEMSELECT

SOLENOID

SYSTEM NO. 2PUMP

SYSTEM NO. 1PUMP

GSE PANELS

F92−064

SYSTEM NO. 1MANIFOLD


PRESSURETRANSDUCER

SYSTEMSELECT

SOLENOID

SYSTEM NO. 2PUMP

SYSTEM NO. 1PUMP

GSE PANELS

F927−023



FLUID LEVELSIGHT GAUGE

FILTER BOWL(PRESSURE)FILTER BOWL

(RETURN)

BLEED VALVE

SAMPLINGVALVE

TEMPERATURESWITCH

Figure 7−10. Hydraulic System Installation




GSEPANEL

F92−065

HANDPUMP

(OPTIONAL)

MANIFOLDRESERVOIR

VARIABLEDELIVERY

PUMP

VARIABLEDELIVERY

PUMP

MANIFOLDRESERVOIR

COLLECTIVE SERVO ACTUATOR

LONGITUDINAL SERVO ACTUATOR

LATERAL SERVO ACTUATOR

DIRECTIONAL SERVO ACTUATOR

SYSTEM 1 SYSTEM 2

GSEPANEL

Figure 7−11. Hydraulic System Block Diagram




Rotor Brake:

A completely separate secondary stand−alone hydraulic system is a part of therotor brake installation. It incorporates a master cylinder operated by the brakehandle in the cockpit, and the actuator that operates the disc brake on the backside of the transmission where the NOTAR� drive shaft connects. A yellowBRAKE caution annunciator in the IIDS secondary display screen warns if thebrake is not fully disengaged.

MASTER CYLINDER WITHINTEGRAL RESERVOIR

HYDRAULICTUBE

CONTROL LINKAGE

F92−066

BRAKECALIPER

Figure 7−12. Rotor Brake System




7−8. PROPULSION SYSTEM

The propulsion system is designed to meet the engine isolation requirements formulti−engine rotorcraft that are defined by the Category A requirements of FARPart 29, paragraph 29.903(b).

Powerplant:

This system consists of two Pratt and Whitney Canada (P&WC) PW207E turbo-shaft engines mounted above the baggage compartment and pointing inboardto drive into the main transmission gearbox (Ref. Figure 7−13 and Figure 7−14).

Each engine is mounted to the fuselage upper deck by a three point, adjustabletitanium mount. The air inlet which is in the middle of the engine is locatedinside a titanium−walled inlet plenum that leads from a flush−mounted inletin the side of the cowling. The combuster end of the engine is surrounded bytitanium firewalls forward, aft, inboard side, and below. It is covered by a fairingdoor, and is ventilated by an exhaust−driven ejector at the aft end of the compart-ment.

PRIMARY EXHAUSTNOZZLE ASSEMBLY SECONDARY

EJECTOR

INSULATIONBLANKET

F92−067

TRIPODMOUNT

TRIPODMOUNT

FWD INLETPANEL

AFT INLETPANEL

ENGINEAIR INLET

REAR STAYASSEMBLY

FORWARDFIRE SEAL AFT

FIRE SEAL

FMUSHROUD

Figure 7−13. PW207E Engine Installation




DCU

T6 THERMOCOUPLE

NP SENSOR

NG SENSOR

OIL PRESSURE PORT

OIL TEMPERATURE PORT

TORQUESENSOR

RH OIL LEVEL SIGHT GLASS

LH OIL LEVEL SIGHT GLASS

CHIP DETECTOR

OIL FILTER IMPENDINGBYPASS INDICATOR

OIL FILTER COVER

FREON PUMP PAD(IF INSTALLED − RH ENGINE ONLY)

STARTERGENERATOR PAD

PMA

FMU

FUEL PUMP

FUEL FILTER

FUEL NOZZLE

AIR INLET SCREEN

FUEL MANIFOLD

F927−057

FMU SHROUD

T6 THERMOCOUPLE

Figure 7−14. Powerplant − Components




Drive system:

A short shaft with flexible diaphragm couplings and anti flail devices connectseach engine to the transmission. A longer shaft with similar couplings drivesthe NOTAR� fan. The main rotor drive shaft connects the planet gear carrierin the top of the transmission to the main rotor hub through a splined connectionat each end. The engines and transmission are electrically bonded to the airframeby suitable jumpers.

BLOWERHOUSINGASSEMBLY

TRANSMISSIONASSEMBLY

STRUTASSEMBLY

STATIC MASTSUPPORT ASSEMBLY

DECK FITTINGASSEMBLY

HYDRAULICPUMP DRIVE

LUBRICATIONPUMP AND FILTER

PRESSURETRANSDUCER

TEMPERATUREPROBE AND

SWITCH

INPUTDRIVE SHAFTS

MAIN ROTORDRIVE SHAFT

F92−069−1

Figure 7−15. Drive System (Sheet 1 of 2)




To minimize structurally−transmitted acoustic noise from the transmission intothe passenger spaces, the transmission is supported from the mast base by eightbolts in elastomeric bushings, and is restrained against rotation by a toothedcoupling arrangement that has a contoured elastomeric ring between the bottomof the mast base and the top of the gearbox.

EXHAUST DUCT

INTERCONNECT DUCT

OIL COOLER

INLET DUCT

AIRFRAMEDECK

VIEW ROTATED

NOTAR FANDRIVE SHAFT

PRESSURE SWITCH(LOW)

MAGNETICCHIP DETECTOR

F92−069−2

Figure 7−15. Drive System (Sheet 2 of 2)

Engine and transmission lubricating oil is cooled by air/oil heat exchangersmounted in the sides of the cowling alongside the transmission. Each cooler is splitso that it serves separately one engine’s requirements plus half of the transmission’srequirements. A direct drive fan on each side of the transmission induces ambientair to flow through the cooler cores. Each engine has its own lubrication pump;the transmission’s pump is located low on the front centerline of the gearbox.

Magnetic chip detectors are provided for each engine and the transmission. Thedetector in the transmission has ‘‘burn−off ’’ capability; the detectors in the enginesdo not.




7−9. ENGINE AIR INTAKE AND INLET PARTICLE SEPARATOR (IPS)

The air intake system provides a path for ambient air to enter each engine compressorcase inlet. The air intake system consists of an inlet screen or optional inlet particleseparator for each engine that prevents debris from entering the engine ducts.

Inlet screen:

The standard inlet screens are 1/4 in. ( 64 mm) steel wire mesh screens locatedon the upper intake cowlings (Ref. Figure 7−16). Each engine inlet screen pre-vents large foreign objects from entering the inlet plenum. A bypass openingis located at the aft end of each inlet screen. The aft facing bypass opening assuresairflow if the screen becomes clogged.

IPS (if installed):

The inlet particle separator is an inertial type particle separator that removesdebris from the ambient air before it enters the engine. The particle separatoris located on the upper intake cowling (Ref. Figure 7−16). Ambient air entersthe particle separator and the air velocity is increased as the air passes overswirl guides. The swirl guides create a vortex that separates heavy particlesfrom the air. The particles drop to the bottom of the particle separator panel.A solenoid valve and bleed air lines route engine compressor bleed air to theparticle separator ejector to eject the particles overboard. The ejector is controlledby the pilot through the IPS switch located on the Utility panel. In the eventthat the particle separator becomes clogged with debris, solenoid operated bypassdoors automatically open for both engines inlets.

NACA inlet:

The NACA engine inlets provide ‘‘ram air’’ for enhanced engine operation/perfor-mance during cruise flight. If the aircraft is equipped with the IPS, the NACAdoors open/close automatically when the airspeed is greater/less than 47 KIAS.A NACA inlet switch is provided on the options switch panel that allows thepilot to override the automatic door opening feature and leave the NACA inletdoors in the closed position. On aircraft with the standard engine inlet screen,the NACA inlet does not include doors, but has a screen covering the inlet. Addi-tional information for operations with the NACA inlet may be found in SectionsII, III, IV, and IX.




F92−070

SOLENOIDSHUTOFF VALVE

PARTICLE SEPARATOR EJECTOR

EJECTORTUBE ASSEMBLY

BLEED AIRTUBE ASSEMBLY

PARTICLESEPARATOR PANEL

BYPASS DOORSOLENOID LATCH

STANDARD INLETSCREEN

BYPASS DOOR

NACA INLET DOOR(IF INSTALLED)

Figure 7−16. Engine Air Intake

7−10.ENGINE POWER MANAGEMENT SYSTEM

Automatic Engine Control:

The Pratt and Whitney PW207E engine is equipped with a single channel FullAuthority Digital Electronic Control (FADEC) which consists of an ElectronicEngine Control (EEC), Fuel Metering Unit (FMU), and fuel pumps. A manualbackup system is provided for emergency operation in case the EEC becomesinoperative. The pilot’s controls for normal operation consist of two rotary enginecontrol switches on the engine control panel for the left and right engines. Theseswitches are gated between OFF and IDLE: the switch knobs must be liftedto pass the gates. The other switch positions are FLY and TRAIN and are notgated. For normal operation, the two twist grips on the collective pitch stickare always left in their NORMAL detent position.

The EEC’s of the two engines are connected together electrically for a torque−matching function, and are both connected electrically to the collective stickand pedal position resolvers for power change anticipation.

When the EEC’s are working properly, the procedure for starting and stoppingrequires no more than selection of the desired engine operation with an enginecontrol switch.

P&WC has built into the PW207E engine the proper shielding to protect theEEC’s from the HIRF threat, and the helicopter’s wiring system components




that are associated with the EEC’s are protected in a similar manner. With thisprotection in place, freedom from lightning damage is also assured.

Train Mode:

Placing an engine control switch in the TRAIN position will simulate a one engineinoperative condition by resetting the selected engine’s governed speed to 92percent NP, thereby putting the engine on standby while allowing single enginetraining on the opposite engine. In the event of an engine failure on the oppositeengine, the engine in TRAIN will automatically revert to 100% NP.

Additionally, the opposite engine will retain the 5 minute Take−off Power engineparameter limiters. The result is more realistic pilot OEI training, providingrotor droop in training if the power requested is above the limiters as wouldhappen in a real OEI condition.

Emergency Manual Control:

The controls for manual operation of the engine power consist of two twist gripson the collective pitch stick and a push button located on the collective control moduleat the end of the collective stick.

The EEC is designed to ‘‘fail−fixed" (EEC’s stepper motor is fixed at its last con-trolled power setting) so there is no sudden change in the level of power if anEEC becomes inoperative. The only noticeable happening is illumination of theyellow EEC/red FAIL warning on the Integrated Instrument Display System(IIDS). No matter at what power level the EEC becomes inoperative, there issufficient travel in the twist grip to control the engine manually from full powerto idle and engine shutdown.

After the EEC becomes inoperative, the pilot uses the appropriate twist gripon the collective stick to modulate the power.




7−11.FUEL SYSTEM

The single crash−resistant elastomeric fuel cell is capable of holding 161.3 U.S.gallons of jet fuel and is located in the lower fuselage under the main cabin floor(Ref. Figure 7−17). It is contained between crash−resistant keelbeams and bulk-heads, with a support panel underneath.

The powerplant separation feature includes a partial−height baffle that runs foreand aft along the bottom center of the cell that provides sufficient fuel reserve forat least twenty minutes of flight following loss of fuel in the other compartment.This provides two separate fuel supplies, and each are capable of transferring fuelfrom the other. This is a pressurized system with a boost pump and jet pump locatedon each side of the longitudinal baffle. With boost pumps operating, fuel is pumpedthrough jet pumps in the opposite fuel cell cavity. The jet pump draws fuel fromthe sump through a pickup and the fuel is ejected on the other side of the longitudinalbaffle.

The fuel system is pressurized having a separate fuel pump located in the sumpin each side of the cell.

The cell is designed with a seven percent expansion space, and has two anti−sloshbaffles across it. Pilot−operated shutoff valves are positioned at the engine firewalls.Self-closing breakaway fittings are installed where fuel lines penetrate the cell wallsand where they penetrate the engine deck. Overboard fuel cell vent lines incorporaterollover valves and flame arrestors located in the vent system stand pipes.

The gravity−type fuel filler port is located on the right side of the fuselage justaft of the pilot’s cockpit door.

Two sump overboard drains for removing sediment and water (one for each sideof the cell) are operated by knobs located under the right side cabin step.

The engine fuel drain system provides a path for residual fuel from the fuel manifoldthat remains after shutdown to be returned to the fuel cell.

A provision is made in the fitting at the aft left hand corner of the cell for makinga connection to an optional auxiliary fuel tank.

The pilot controls the fuel system by the Fuel System Panel mounted switches.Fuel level is sensed by a forward probe and an aft probe, and is displayed on theIIDS. Two fuel pressure switches activate caution lights in the IIDS when the pres-sure falls below the acceptable limit.




FUEL METERING UNIT (FMU)SHROUD BOX (2 PL)

ENGINE DECKFUEL SHUTOFF

VALVE (2 PL)

FUEL PRESSURESWITCH (2 PL)

RH FUEL FEEDVAPOR SHROUD

TEE FITTING (2 PL)

RH FUEL FEED SYS

TO FWD RHVENT SYS

FUEL FILLERASSY

FLAMEARRESTOR (2 PL)

VENT OVBDDRAIN (2 PL)

SUMP DRAIN CABLE

SPRING LOADEDFLAPPER VALVE

FUEL BOOSTPUMP (2 PL)

SUMP OVBDDRAIN (2 PL)

FUEL QTY PROBE(2 PL) LEFT AND RIGHT

SUMP DRAINVALVE (2 PL) FUEL CELL

LOW FUEL LEVELSENSOR (2 PL)

CENTER BAFFLE

CENTER BAFFLEFUEL BOOSTPUMP (2 PL)

FEED LINE TO R ENGINE FEED LINE TO L ENGINE

"T" FITTING (2 PL)

R FUEL XFER TUBINGL FUEL XFER TUBING

EJECTOR PUMP (2 PL)

CHECK VALVE (2 PL)

FUEL TRANSFER SYSTEM

FUEL CELL OUTLETVAPOR SHROUD

FWD LHVENT SYS

LH FUEL FEEDVAPOR SHROUD

LH FUELFEED SYS

FUEL FEED FRANGIBLECONNECTOR (2 PL)

ENGINE DECK FRANGIBLECONNECTOR (2 PL)

TEE FITTING VAPORSHROUD DRAIN (2 PL) TOENGINE DRAIN SYSTEM

VENT−ROLLOVERVALVE (4 PL)

TO AFT LHVENT SYS

LEFTENGINE

RIGHTENGINE

LOOKING DOWN

GRAVITY FILL VALVEAND FLAPPER VALVE

F92−073

Figure 7−17. Fuel System Schematic




NOTE: If the voltage for the probe drops below the specified operating limit, thesegments in the fuel quantity vertical scale blank with the digital quantity stillactive.

FUEL

FUEL LOW CAUTIONSEGMENTS (YELLOW)

FUEl QUANTITYSEGMENTS (GREEN)

FUEL LOW WARNINGSEGMENT (RED)

LB

FUEL LOW WARNING TICK MARK (RED)

FUEL FLOW LINELOW FUEL PRESSURE

ANNUNCIATORS (YELLOW)

FUEL FILTER IMPENDINGBYPASS ANNUNCIATOR (YELLOW)

FUEL SHUTOFF VALVE POSITIONANNUNCIATOR (YELLOW)

LOW FUEL PRESSUREANNUNCIATORS (YELLOW)

CURRENT FUEL QUANTITYDIGITAL DISPLAY (WHITE)

FUEL FILTER IMPENDINGBYPASS ANNUNCIATOR (YELLOW)

F92−072

Figure 7−18. IIDS Fuel System DisplayFuel quantity (FUEL) is shown by a vertical bargraph inside a fuel tank iconrectangle, with the corresponding digit value in pounds, shown immediately below.The green bar shortens vertically from the top as fuel is burned proportional tothe total tank volume. When the green box disappears, two yellow segments illumi-nate below to indicate a low fuel caution (approximately 45 minute reserve). Whenthe last yellow segment disappears, a red segment illuminates below to indicatelow fuel (approximately 20 minute reserve). Independent left and right fuel lowwarning red ‘‘tick’’ marks beside the red segments are activated when the low levelsensor reaches the warning level of 100 lbs.

Fuel flow to the engines is shown below the fuel quantity bargraph: Connectionsfrom the fuel tank to each engine is shown immediately below the digit value offuel quantity. A solid line indicates normal fuel flow and alternating white and yellowoffset segments indicate low fuel pressure.

The display of fuel valve left and right engine position is shown by a segmentabove and below each fuel line for the respective left and right fuel valves. During




the time a fuel valve is in transit between open and closed positions, the fuel valveindications will flash. Fuel valve in transit is defined by both fuel valve input discreetsbeing open circuit.

The fuel filter impending bypass status is shown by an inverted ‘‘U’’ above eachfuel line indication.

7−12.FIRE EXTINGUISHING SYSTEM

The fire extinguishing system provides a means for the pilot to direct a chargeof fire extinguishing agent into the designated fire zone of each engine. There isno fire extinguishing system for the transmission area.

Refer to Section III, paragraph for fire emergencies.

The fire extinguishing system (Ref. Figure 7−19) contains two individual hermetical-ly sealed pressurized spherical containers (bottles) that are filled with 60 cubicinches (16.38 CC) of CF3 BR (Bromotrifluoromethane), also known as Halon 1301,and pressurized with nitrogen gas to an internal pressure of 700 PSIG (49.22 kg/cm2).Each bottle serves as the primary bottle for its appropriate side engine.

Each bottle is equipped with dual outlet ports, a pressure gauge with electricallow pressure warning signal to IIDS, filler port and thermal relief valve. The outletsports are fitted with electrically discharged explosive squibs. The fire extinguishercartridges are armed and ready for firing when the fuel shutoff valves are placedin the OFF (closed) position. The bottles are discharged when the BOTTLE DIS-CHARGE switch is momentarily placed in the PRI (primary) or ALT (alternate)position.

The BOTTLE DISCHARGE switch is a momentary type, three position switch lo-cated between the left and right fuel shutoff valves on the cockpit FUEL SYSTEMpanel. Placing a fuel shutoff valve OFF arms the fire extinguishing system for thatengine and selection of PRI discharges its primary bottle. Selection of ALT dischargesthe second bottle onto the same engine.




DISTRIBUTION TUBERIGHT SIDE

DISTRIBUTION TUBELEFT SIDE

REFENGINEDECK

REFENGINE DECK

DISCHARGE TUBE

OUTLETPORT

DISCHARGETUBE

OUTLETPORT

CROSS FLOWTUBES

FIRE BOTTLEASSEMBLY

(BLUE PORT)ALTERNATECARTRIDGE

(RED PORT)PRIMARY

CARTRIDGE

PRESSURE GAUGE

FILLER PORT

F92−146

FUEL SYSTEML BOOST R BOOST

ON

OFF

BOTTLE

PRI

ALT

LEFT OFF RIGHT OFF

ON

OFF

FUEL SHUTOFF

DISCHARGE

OFF

FIRE EXTINGUISHERBOTTLE DISCHARGESWITCH

Figure 7−19. Fire Extinguishing System




7−13.ELECTRICAL SYSTEM

The electric system is designed to maintain separation of the power generating sys-tems. Wiring for each system is physically separated to each side of the helicopter tothe greatest extent possible. Power from the two generators does not pass togetherthrough a single connector at any point on the aircraft to preclude any single pointfailure that could result in loss of power to the essential bus.

F92−075B

LEFTSTARTER

GENERATOR

LEFT

START RELAY

LEFT BUS TIE RELAYLEFT PWR RLY

LEFT GEN BUS

EXTERNAL POWER

LEFT AVNCS BUS

LEFT AVIONICS RELAY BAT

RIGHTSTARTER

GENERATOR

RIGHTSTART RELAY

LEFT BUS TIE RELAYRIGHT PWR RLY

RIGHT GEN BUS

RIGHT AVNCS BUS

RIGHT GCU

RIGHT AVIONICS

RELAY

SHUNT 2

LEFT GCU

SHUNT 1

ESS PWR

ESS PWR

BATTERY RELAY

RT ESS BUSRELAY

RIGHT ESS BUS

LEFT ESS BUS

BATTERY BUS

EXTERNALPOWERRELAY

RIGHT DC BUS

LEFT DC BUSRT

LT

LT ESS BUSRELAY

Figure 7−20. Battery Power and External Power Subsystem Block Diagram

Two engine−mounted starter−generators rated at 200 amperes each provide29 volts DC to the aircraft. Bus tie relays provide redundancy by allowing eithergenerator to provide power to all busses.

When installed, a generator cooling option allows aircraft operations in higher ambi-ent temperature conditions (Ref. Figure 7−21).

The left and right essential bus relays allow the left and right essential busesto be powered by either of the two generators, or by the battery if all power fromthe generators is lost.

Starter and generator functions are directed by individual generator controlunits (GCU), each of which provides starter control, voltage regulation, and protec-




tive functions. Electric power is distributed by two electric busses and a batterybus. A starter contactor connects the starter generator to the battery bus. Aftera successful start, the starter−generator begins generating current and is broughton line by the GCU through the generator contactor.

The pilot monitors generator load on the IIDS. The pilot can manually reset or dese-lect either generator by using the generator switches located on the Electrical Masterpanel.A standard 22 ampere−hour nickel−cadmium battery is used for engine startand for reserve electric power. A 27 amp hour and 44 amp hour (aft mounted batteryonly) are also available as options. The battery relay and external power relay arecontrolled by the power switch on the Electrical Master panel. The standard mount-ing of the battery is in the nose of the helicopter, however, an aft−mounted batteryis available as an option.

BATTERY

POWER AND L/R GENERATOR SWITCHES

EXTERNAL POWER BOX RELAY(FRONT MOUNTED BATTERY)

EXTERNAL POWERRECEPTACLE

STARTER/GENSTARTER/GEN

ELECTRICALLOAD CENTER

GENERATOR CONTROL UNIT

GENERATOR CONTROL UNIT

F927−091A

OPTIONAL REAR−MOUNTEDBATTERY

EXTERNAL POWER BOX RELAYFOR REAR−MOUNTED BATTERY

VIEW LKG INBD

AIR VENT INLET (REF)

GENERATOR COOLING INLET (IF INSTALLED)(RH SHOWN; LH OPPOSITE)

GENERATOR COOLING INLET

Figure 7−21. Battery Power, External Power, and DC Power Component Locator




The key switch is located on the right hand side of the instrument panel. Allswitches and brightness controls that operate the electric system are on the console.

The ground power receptacle for 28 volts DC is in the right hand side of thefuselage below and forward of the pilot’s door.

Two grounding jacks are located on the right hand side of the fuselage, one ad-jacent to the ground power receptacle (forward mounted batery) and one adjacentto the fuel filler port.

Circuit breakers for essential circuits are located in the cockpit on the Left andRight Essential Bus panels; nonessential breakers are located in the baggagecompartment ceiling. One 29 volt DC outlet is located in the cockpit on the copilot’sside of the console, and another one is on the left hand cabin wall aft of the cabindoor.

Aircraft Lighting:

Aircraft Interior Lighting:

Cockpit:

Floodlight (1)

Map Light (1)

Instrument Floodlights (3) (Powered By Right Essential Bus)

Main Cabin:

Threshold Lights (2)

Baggage Compartment:

Floodlight (1)

Aircraft Exterior Lighting:

Nose:

Fixed Landing Light (1)

Fixed Hover Light (1)

Empennage:

Left End of Horizontal Stabilizer:

Red Navigation Light (1)

Right End of Horizontal Stabilizer:

Green Navigation Light (1)

Top Center of Stabilizer:

Flashing Red Anticollision Light (1)

White Navigation Light (1)

Bottom of Tailboom, Forward of Thruster:

Flashing Red Anticollision Light (1)

White Navigation Light (1)




7−14.ENVIRONMENTAL CONTROL

The environmental control system for the helicopter consists of the ventilation sys-tem and the heat/defog system.

Ventilation System:

Ambient air is taken in through an inlet in the right side of the upper cowling,is directed through a water separator and a two−speed fan, and into a manifoldthat distributes the air to the cockpit and to the cabin − then out of a port inthe baggage compartment door. In the cockpit, four adjustable gaspers, two onthe windshield’s center bow blow outboard toward the pilots’ heads, and twoon the forward door frame blow inboard toward their lower torsos. Six adjustablegaspers are mounted in the ceiling of the cabin. The fan speed switch is locatedon the Utility Panel.

Secondary ventilation for the cockpit is provided by two conventional clear plasticadjustable snap vents in the window of each cockpit door.

Heat/Defog System:

The heat source is bleed air from the compressors of the two engines. This hotair is directed through a pilot−operated on/off valve located on the Utility panelto a pair of ejectors that mix bleed air and ambient air to a desired temperatureand flow rate. One ejector serves the cockpit; the other serves the cabin.

The cabin ejector is located low on the right side of the cabin just aft of thedoor. Its discharge air is directed across the cabin under the rear seats. An adjust-ing lever for controlling the bleed air admitted to the ejector, and so the dischargevolume, is recessed in the wall at head height directly above the ejector.

The cockpit ejector is located in the compartment below the pilot’s seat, andis operated by a push/pull control mounted vertically along the right hand sideof the console. From the ejector, warm air is ducted forward to two aft−facingnozzles above and forward of the pilots’ feet, and to a pair of nozzles along thebottom of the upper windshield panels to defog them. Each pilot has a push/pullknob located under the instrument panel to operate a butterfly valve that modu-lates the airflow toward his/her feet.

An automatic disconnect monitored by the IIDS cuts off all bleed air whenevereither engine becomes inoperative in flight to maximize the operating engine’spower output to the rotor. This cutoff function maybe overridden by placing theCAB HEAT switch in the OVRD position.




CREW STATION

PASSENGER COMPARTMENT

PASSENGERCOMPARTMENTEJECTORS

ENGINEENGINE

CHECKVALVES

START−UPLOCK OUT

HEAT ON/OFFSWITCH

DEFOGGINGMANIFOLDS

FOOT HEATERS FOOT HEATERSCONTROL VALVES

FLOW CONTROLVALVES

FLOWCONTROL

VALVES

FLOW CONTROLSHUT−OFF VALVE

(ON/OFF)

PILOT HEATEJECTORS

F92−076

Figure 7−22. Heat/Defog System Schematic




7−15.INTEGRATED INSTRUMENTATION DISPLAY SYSTEM (IIDS)

The IIDS provides for the monitoring and display of various aircraftparameters and for caution/warning annunciation. The baseline config-uration includes a set of engine, drive train, rotor, NOTAR�, electrical,fuel, hydraulic, and caution/warning indicators. It also incorporates abuilt−in rotor and NOTAR� fan balance system and stores system oper-ating and exceedance parameters for enhanced maintainability.

The IIDS accepts analog and discrete inputs from various aircraft sub-system transducers and provides signal conditioning and conversionto digital format. Once converted to digital format, this informationis provided to the display electronics for the cockpit display and to aserial port for access by a data recorder or computer. Also, limit checkingon certain parameters is performed to provide the caution/warning an-nunciation. The display is a color, Liquid Crystal Display (LCD) panelwhich allows the flexibility of integrating the specified sensor data andcaution/warning information onto a display packaged as one unit.

Three levels of Built−in−Test (BIT) are used to determine system health,including Power−up, Continuous, and Commanded BIT. Power−up diag-nostics will check the health of each function or module within the IIDSand display this test status. Continuous testing checks the operationof the IIDS during aircraft operation and displays and/or logs any fail-ures. Commanded BIT, initiated using the IIDS keyboard, performsa display test, along with those tests performed during Continuous BIT.The display is put into ‘‘lamp test’’ mode, where all segments are acti-vated, so that the display can be visually inspected for segment failures.Both Power−up and Commanded BIT test the two engine and the trans-mission fire detectors, and the bleed air leak detector (if installed).

General

Built−In−Test




F92−077

ROTORSYSTEM

HYDRAULICSYSTEM

AIRFRAMESYSTEM

POWERPLANTSYSTEM

BALANCEMONITORING

SYSTEM

INTEGRATED INSTRUMENTDISPLAY SYSTEM (IIDS)

DRIVESYSTEM

FUELSYSTEM

ELECTRICALSYSTEM

NOTAR�SYSTEM

Figure 7−23. IIDS System Monitoring

BIT failures are stored in non−volatile memory to assist in three situa-tions:

First, a transient or intermittent failure;

Second, a situation where the pilot observed a problem with the IIDSbut didn’t notice any failure annunciation;

Third, ease of IIDS maintenance on and off the aircraft. These faultwords are stored in the Fault Log when a BIT failure was detectedin the IIDS, BMS, EEC, or aircraft transducers/sensors, and can beexamined through the IIDS display or ground based maintenance com-puter (GBMC).

When the testing determines that an internal fault exists, the appropri-ate redundant function, if such redundant system exists, will be com-manded to assume the primary role. The redundant functions shallbe sufficiently isolated such that a failure of the function will not cause

BITFailures




the failure of another function. These fault monitoring provisions areimplemented using hardware Built−in−Test Equipment (BITE) and soft-ware diagnostics, allowing isolation of failures to at least the internalmodule level. In addition, provisions are made to check operation ofthe transducers and sensors and provide an appropriate maintenancealert.

Any sensor that can be checked for proper function and is determinedto have failed causes blanking of the digit display for that parameter.

The following are exceptions to the above:

1. A failure of a sensor for the primary display parameters (EGT,Torque, NR, and NP) causes both the vertical scale and digit valueto blank.

2. If the voltage for the fuel probes (Battery Bus voltage input) dropsbelow the specified operating low limit of 18 volts for more than40 seconds, the low voltage indication shall be to blank the segmentsin the fuel quantity vertical scale. The digit quantity shall remainactive. When the probe voltage goes back above 18 volts for morethan one second, the vertical scales shall be illuminated. A failureof one of the fuel probes causes only the digit values to blank whereasthe failure of both probes causes both the vertical scale and digitvalue to blank.

3. If the parameters displayed on the alphanumeric display (Pressure/Density Altitude, L/R engine fuel flow, CLP, and Hydraulic Pressure)are out−of−range, the display will read NOT VALD.

The functional architecture of the IIDS to meet these design goals andthe operational requirements of the aircraft is shown in Figure 7−23.

SYSTEM OPERATING PROGRAM: The System Operating Programprovides the programming and functions controlling the data collection,displays and formatting, key entry functions, date/time clock, cautionsand warnings, exceedance detection, memories and BIT feature.

SystemSoftware

Architecture




The integrated display is divided into two separate displays. The prima-ry is on the right side and the secondary is on the left.The primary display includes the following information.

Power turbine speed: NP

Rotor speed: NR

Measured gas temperature: EGTEngine Torque: TORQUE

NR and NP are displayed with three vertical bargraphs and a digitvalue of NR displayed in the center.

Torque display: Displayed in % torque. The IIDS obtains engine torquefrom the EEC. If the EEC fails, the IIDS calculates torque by usingNG, OAT, and pressure altitude measurements. The vertical bargraphsand three digit indicators on this display indicate torque in percent(%). The vertical bar has four ranges as defined below:

Green segments indicate continuous operating range including Maxi-mum Continuous Power (MCP).

Yellow segments indicate:

Transient Take Off Power (TOP) operating range (5 minute limit)

OEI operating range (2.5 minutes)

Top red segments − do not exceed limit.

NOTE: Even though the IIDS displays engine torque, the transmission sets the torquelimit for helicopter operations, and therefore, the displayed torque limits arelower than those for the engine as stated in the Pratt & Whitney MaintenanceManual.

EGT is indicated by two vertical bargraphs and a three digit indicatorshowing EGT in 1°C increments. Displayed on the IIDS as EGT. Pratt& Whitney refers to this measurement as MGT (Measured Gas Tempera-ture). The vertical bargraph has four ranges as defined by the displaymode; they are:

Green segments: continuous operating range (MCP operating range)Yellow segments: transient operating range (TOP or OEI)Top red segment: do not exceed limit

NOTE: The IIDS provides a time count−down on the alphanumeric display when thepilot enters TOP, OEI, or transient flight conditions. Should the pilot exceedthe count−down, the IIDS then provides an time overcount, and exceedanceand data logs are created.

Warning annunciators in red are for EGT, Torque, NR, NP, and EEC FAIL.Caution annunciators in yellow are for EEC minor fault, EEC MAN(manual) mode, and OEI (one engine inoperative).

Primary andSecondary

Displays




In the secondary display, caution annunciators in yellow are givenfor engine chips, engine oil temperature, high or low, engine oil pressurehigh or low, generator load high, generator out, NG high, transmissionchips, transmission oil temperature high or low, fuel pressure low, fuelfilter impending by−pass, fuel valve closed, battery warm, rotor brake,hydraulic system status, baggage door open and IIDS status.

Engine oil pressure display: Displayed in % PSI, and is a function ofNG speed and engine oil temperature.

Engine oil temperature display: °CGas producer turbine speed display: %NG

Transmission oil pressure display: Displayed in % PSI, and is a functionof NR speed and transmission oil pressure.

Transmission oil temperature display: °CGenerator load display: %LOAD

Warning Annunciators displayed in red are shown for engine fire, engineoil temperature high, engine oil pressure high or low, NG high or low,transmission area fire, transmission oil pressure high or low, fuel quanti-ty low CAB HEAT (bleed air leak), and BAT HOT.

Day or night modes may be selected using the Light Master switchlocated on the Lighting Control Panel. Placing the Light Master switchON selects night mode.

The display brightness is adjustable using the inner ring of the IIDScontrol potentiometer also located on the Lighting Control Panel.

When in the night mode, the IIDS will automatically increase displaybrightness when a caution/warning message is received and displayed.To return to the preset brightness, press the CLR key momentarily.

LIGHTING CONTROL

LT MSTR CONSOLE IIDS

FLOOD INSTR

STROBE AREAPOSN

ON

OFF OFF OFF

ON

OFF

ON

OFF OFF

BOTH

CKP

CAB

IIDS DISPLAYBRIGHTNESSCONTROL

F92−078

LIGHT MASTERSWITCH

Figure 7−24. IIDS Display Brightness Control

The IIDS has a 2 line by 16 character alphanumeric display. This displayallows messages to be displayed regarding systems limit exceedance,

Display Brightness

Controls

AlphanumericDisplay




condition, various cautions and warnings as well as expanded featuresof the IIDS to be viewed by the pilot. Yellow and red segments are locatedto the left of each line that indicate if the associated message is a cautionor a warning. The expanded features of the IIDS are selected in conjunc-tion with the IIDS keyboard.

Certain conditions will cause the alphanumeric display to automaticallydisplay a message.

At start−up, and if required during flight, messages are displayed onthe alphanumeric display automatically. A list of these messages isfound in Table 7−1. This table also defines the priority of the messageto be displayed, the classification of the message (warning/caution/advi-sory W/C/A), and whether the message can be cleared (CLR) from thedisplay.

The IIDS uses the following logic to determine an aircraft on−ground/off−ground condition. The IIDS uses this information to enable or disablecertain caution/warning, indications and alphanumeric display advisorymessages.

Aircraft on−ground if:1. NR � 80%

ORAircraft on−ground if:

1. NR >80%, and2. CLP <5%, and3. Torque (either engine) >10%

Otherwise, the aircraft is off−ground. The transition from one conditionto the other is not recognized until after the new condition has existedfor 5 seconds.

AOGLogic




NOTE: Advisory messages may not indicate a malfunction or emergency.


NPNR

NPENGOUT

TORQUE EGT

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

F92−079

WARNINGANNUNCIATOR (RED)

ALPHA−NUMERICMESSAGEDISPLAYWARNING

ANNUNCIATOR (RED)


Figure 7−25. Alphanumeric Display

Table 7−1. Automatic Alphanumeric Display Warning/Caution/Advisory Messages

SAMPLE MESSAGE(Fault)

CAUSE FOR DISPLAY W/C/A CLR CORRECTIVE ACTION

ENG POWER CHECKL PA CK NG −1.8LPA CK EGT−11.2

Invalid performancemargin (power checkfailed)

W YES Advise Maintenance

OVR TORQ LFT 2:30OVR TORQ RT 2:30OVR EGT LFT 2:30OVR EGT RT 2:30OVR NG LFT 2:30OVR NG RT 2:30

MTO or OEI overcount W NO Advise Maintenance

PRES 1 = 0 PSIPRES 2 = 1000 PSI

Hydraulic system sta-tus: activated on hy-draulic caution indica-tion1

C YESPerform malfunctionprocedure. Ref. Section III.

TEMPERATUREHydraulic system over-temperature

C YESPerform malfunctionprocedure. Ref. Section III.

L ENG OIL COLDEngine oil temperaturecold (Starting)

C NO

Start engine with enginecontrol in IDLE. Do notadvance to FLY untilmessage blanks

RIGHT STAB FAIL

LEFT STAB FAIL

TOTAL STAB FAIL

Right Stabilizer Actuator Failure

Left Stabilizer ActuatorFailure

Both StabilizerActuator Failure

C

C

C

YES

YES

YES

Perform malfunctionprocedure. Ref. Section III.

CHK BMS SENSOR BMS sensor(s) failure2 C YES Advise maintenance




Table 7−1. Automatic Alphanumeric Display Warning/Caution/Advisory Messages

SAMPLE MESSAGE(Fault) CORRECTIVE ACTIONCLRW/C/ACAUSE FOR

DISPLAY

TQ SPLIT EXCEEDTransmission inputtorque split exceedance

C YES Ref. Section III

TORQ LFT 2:30TORQ RT 2:30EGT LFT 2:30EGT RT 2:30NG LFT 2 :30NG RT 2:30

TOP or OEI countdown A NO N/A

CARGO HOOK OPEN Cargo hook open A YES N/A

IPS BYPASSParticle separatorclogged: IPS in bypass

A YES Ref. Section III

NACA DOORNACA door in the incor-rect position.

A YES Advise Maintenance

BATT DISCHARGE Battery Discharging A YES N/A

EXTGSHR PRESS LOOptional fire extin-guishing system bottlepressure low


RECORDING DATACrew commanded datarecord

A NO N/A

DATA XFER CMPLDownload of data logsto ground based com-puters

A YES N/A

CHECK FAULT LOGFault Log duringflight2,3 A YES Advise Maintenance

CHK EXCEED LOGExceedance Log duringflight3 A YES Advise Maintenance

CHECK ROTOR BALMain rotor out of bal-ance


CHECK NOTAR BALNOTAR fan out of bal-ance


NOTE: 1.With a single system failure, this message is displayed when the failed systempressure decreases to 250 PSI. This message will reappear when remainingoperating system pressure decreases to 500 PSI.

2.This caution does not affect dispatchability.3.These messages are generated for conditions that create a fault log or an

exceedance log and are displayed only when the aircraft is on−ground as determined by the AOG logic.




7−16.IIDS DATA STORAGE

Selected information that is monitored by the IIDS for display is alsostored by the IIDS. Information is stored in non−volatile memory andis available by selecting various menu functions through the front panelkeys. All information is available using a ground based maintenancecomputer (GBMC). A partial Data Log and Cumulative Log are viewableon the alphanumeric display. Complete Data and Cumulative logs areavailable through the GBMC. The operating data is stored in one ofsix data records:

LOG TYPE A/N DISPLAY

Data Log NO

Cumulative Log NO

Exceedance Log YES

Fault Log YES

Trend Log YES

Setup Log YES

The data log provides one and a half minutes of data collection. Thedata is recorded in a continuous memory buffer ‘‘loop’’ and will be contin-uously over−written unless an exceedance occurs, or the crew requestsa record of an event. Exceedances generate both data logs and excee-dance logs. The data log record provides a ‘‘window in time’’ to examineevents around an exceedance or other incident. The data log can storefive of these events. This information is accessed by the GBMC only.

Pilots may generate a data log by first clearing the alphanumeric display,then pressing and holding the �REC" button for 7 to 10 seconds beforereleasing. The message RECORDING DATA will then be displayed onthe alphanumeric display.

The cumulative log retains data concerning the aircraft operational his-tory and current configuration. As the aircraft configuration changes(e.g. component changes) the Cumulative Log will be updated by themaintainer on the GBMC. The IIDS will only retain one CumulativeLog in memory.

NOTE: Time Summary Menu procedure (Ref. Figure 7−28) may be used to accesscumulative usage data as well as flight time data.

Data Storage

Data Logs

Cumulative Log




The exceedance Log provides a ‘‘snapshot’’ record of the parameter dataat a particular moment in time. This type of record occurs whenevera parameter exceedance is detected. This recording function is onlyactive when NG for either engine AND NR is greater than 50 % ANDEGT on either engine is greater than 400°C

This feature is capable of storing 100 exceedance Logs. Exceedancesare recorded for EGT, engine torque, transmission torque, NG, NP, NR

(high), engine oil pressure (low), transmission oil pressure (low), andcargo hook.

Exceedance logs maybe either recorded or downloaded to the GBMC.

Perform Exceedance Log Menu procedure (Ref. Figure 7−41) to accesscumulative usage data.

The Fault Log contains data associated with fault discrete data fromthe EEC’s and a BIT failure that was detected in the IIDS, BMS, oraircraft transducers/sensors. This type of log is recorded whenever anIIDS, EEC, or aircraft sensor fault is detected. The system is capableof storing 100 Fault Logs.

Fault logs maybe either recorded or downloaded to the GBMC.

Perform the Fault Log Menu procedure to access data (Ref. Figure 7−43).

IIDS setup contains (Ref. Figure 7−44) data that reflects the currentconfiguration of the aircraft, such as, aircraft serial number, enginetype installed, engine serial number, fuel calibration, operating soft-ware, etc.

On power up, the IIDS uses information in the setup log to comparethe current Torque and EGT trim values from the EEC to the valuesstored in the IIDS setup log to assure the data collected by the IIDSremains with the respective engine. If there is a discrepancy, a faultlog is created and certain engine ASCM functions are disabled for theaffected engine(s): Exceedance Logs for NG, NP, Torque, EGT; PowerAssurance function (including trend logs); and Cumulative Logs (cyclecounting, SSO, FSO, TSN, and engine run time).

To recover from this disable function, the Setup Log must be revisedto match the values from that specific engine(s) through the GBMC.Once the Setup Log has been revised, a power−up of the IIDS will verifythe new data. If the new values match, all engine ASCM functions arerestored.

ExceedanceLog

FaultLog

IIDS Setup Log




The IIDS features 7 keys on the right side of the front face to allowthe pilot access to the various functions/programs by paging throughthe menus. The keys include:

‘‘CLR’’ (clear): Used to blank the alphanumeric display and exit allmenu functions if pressed for more than 1.5 seconds. If pressed for lessthan 1.5 seconds in the Night Mode after a C/W/A event, the CLR keyresets the intensity to the previous setting.

MENU: Used to access the next higher level of the menu structureor to enter the top level menu from display blanked and to return tothe ‘‘action’’ menu with edit fields not updated.

UP ARROW � : Used to scroll between menu and submenu names,or between data and message items. Holding this key for more than2 seconds initiates automatic scrolling, at approximately one item persecond. When the scrolling reaches the end of the menu the scrollingfeature loops back to the start of the menu.

DN (down) � ARROW: Same as the UP ARROW, except scrolls inthe opposite direction.

ENT (enter): Used to enter a menu or submenu after it has been selectedwith the ‘‘�’’ or ‘‘�’’ keys, enter an ‘‘Action’’ field within a menu selectionthat is bracketed by ‘‘< >’’ to allow editing, and to advance to the nextedit digit (or field within the ‘‘Action’’ field. The digit (or field) that canbe edited will flash.

‘‘REC’’ (record): Used to initiate crew requested Data Log and to enterinto memory data that is used to initialize the TIME/DATE, ENGINEPARM, and Cargo Hook CALIB CODE and FUEL CALIBRATION func-tions in the IIDS. When the key is pressed for 7 to 10 seconds, theparameter data from 45 seconds prior to and 45 seconds after key activa-tion, is stored in nonvolatile memory. The message RECORDING DATAis displayed on the alphanumeric display during this time.

‘‘DISP’’: Used to change the display from ‘‘display by exception’’ to ‘‘con-tinuous display’’ when the key is pressed for less than 1.5 seconds. Inthe exception mode, the secondary display screen area is blank unlessone of the limits is within a predetermined range of it’s caution limitvalue. When this happens, the digit display of the particular limit willrevert to continuous display until the parameter value drops below thepredetermined threshold. If the exceedance parameter enters cautionor warning range the appropriate caution or warning displays are illumi-nated.

When the ‘‘DISP’’ key is held for more than 1.5 seconds the IIDS performsa BIT test and the front panel display will show all LCD segments ina lamp test mode.

CLR

MENU

ENT

REC

DISP

J1

F92−080

IIDSKeyboard




7−17.BALANCE MONITORING SYSTEM

NOTE: Helicopter gross weight should be at 5200 �300 LBS before performing ‘‘MainRotor Balance’’ procedure.

The BMS program is an integrated vibration monitoring system whichcalculates and displays balance solutions for both main rotor and NO-TAR� fan. The intention of this integrated balance system is to eliminatethe requirement to fly dedicated tracking/balance flights.

The system is linked to three vibration sensors on the airframe andtwo position pickups on the main rotor and the NOTAR� fan. The stan-dard BMS program is a ‘‘smart chart’’ system. For most balancing actionsthe user will simply follow the directions of the BMS Alpha−numericdisplay (Ref. Figure 7−29 thru Figure 7−40). The normal sequence ofevents is for the pilot to request the BMS program from the IIDS bypressing the ‘‘MENU’’ key on the IIDS panel and paging down the menuto BMS. The BMS system will then analyze the input from the rotor/fansensors and calculate a correction and display this information in theIIDS Alpha−numeric display.

ÎÎÎÎ

1. IIDS2. BMS SIGNAL PROCESSING UNIT3. BMS SENSOR CABLE HARNESS4. AZIMUTH SENSOR (MAG PICKUP/PHOTOCELL)5. VIBRATION SENSOR (VELOCIMETER) F92−081

Figure 7−26. Balance Monitoring System Installation

An optional item to the BMS is a Spectrum Analyzer Vibralog. Thesoftware for this program resides within the GBMC. Spectrum analysisallows downloading to the GBMC and viewing of the entire vibrationspectrum of the rotor and the NOTAR� fan. The system allows theoperator to analyze vibrations, other than rotor/fan, and determine theprobable source by comparison with known component frequencies.

Standard BMSProgram

Optional Spectrum

Analyzer




7−18.IIDS MENU STRUCTURES

F92−082

For expanded menu structure

Ref. Figure 7−46

ELAPSED TIME

MM.SS

TOP LEVEL

TIME SUMMARY

POWER CHECK

BALANCE MONITOR

AIRCRAFT MONITOR

CLPXXX PERCENT

PRES ALT XXXXFT

DENS ALT XXXXFT

TIME/DATE

FUEL CALIBRATION

HOOK WT XXXX LBS

ENT

ENT

ENT

ENT

Continuous display of collective position

Continuous display of altitude

‘‘ENT’’ Key resets, starts and stops timer (alternate action);

‘‘CLR’’ Key exits function and resets timer


Ref. Figure 7−29 thru Figure 7−40


Ref. Figure 7−41 thru Figure 7−44


Ref. Figure 7−45

SET ENGINE PARMENT


Ref. Section VENT


Ref. Figure 7−28ENT

ENTFor expanded menu structure

Ref. Section X


Ref. Figure 7−47

L ENG WF XXX PPH

L ENG WF XXX PPHContinuous display of fuel flow

Figure 7−27. IIDS Top Level Menus




TIME SUMMARY

GEAR BOX

TSO =

RT ENGINE TIME

ENT

LST FLT TIME

ENT

NOTE 1: THIRD LEVEL MENU FOR RIGHT ENGINE

SAME AS FOR LEFT ENGINE

POWER MODULE

TSO=

CMPSR TURB CYCLE

CNT ACCUM=

POWER TURB CYCLE

CNT ACCUM=

NOTE: TO RETURN TO PREVIOUS HIGHER LEVEL − PRESS MENUF92−083

TOP LEVEL SECOND LEVEL THIRD LEVEL

TOT FLT HR

TOT FLIGHTS

LFT ENGINE TIME

IMPELLER CYCLE

CNT ACCUM =

NOTE 1

Figure 7−28. Time Summary




ACQUISITION

COMPLETE

MAIN ROTOR TRACK

MAIN ROTOR

SOLUTION OPTIONS

RUN XX M/R

MEASUREMENTS

BALANCE MONITOR COLLECT M/R DATAMAIN ROTOR

BALANCE

F92−084NOTE: TO RETURN TO PREVIOUS HIGHER LEVEL − PRESS MENU

FLY 100% GROUND

PRESS REC

ACQUIRING

100% GND LAT

FLY HOVER IGE

PRESS REC

FLY 120 KIAS

PRESS REC

ACQUIRING

100% GND LAT

X.XX IPS AT YY.YY

NOTAR

BALANCE

SPECTRUM

BMS FAULT LOG

BMS VERSION LOG

MAIN ROTOR

CONFIGURATION


DISPLAY M/R

SOLUTION RUN XX

NOTE 2

NOTE 3

ACQUIRING

HOVER IGE LAT

HOVER IGE LAT

X.XX IPS AT YY.YY

NOTE 3

ACQUIRING

120 KIAS LAT

NOTE 2

120 KIAS LAT

X.XX IPS AT YY.YY

NOTE 3

NOTE 2

120 KIAS VERT

X.XX IPS AT YY.YY

NOTE 4NOTE 2: AUTOMATICALLY STEPS THROUGH ACQUIRING

MEASUREMENTS SPECIFIED FOR THIS REGIME.

NOTE 3: WHEN COMPLETE, THE RESULT IS DISPLAYED

FOR 4 SECONDS AND THE DISPLAY GOES TO

NEXT REGIME.

NOTE 1: WHEN COMPLETED, MESSAGE IS DISPLAYED

FOR 1 SECOND

RUN XX RPM XXX

120 KIAS VERT

FLY 80 KIAS

PRESS REC

ACQUIRING

ACQUIRING

80 KIAS LAT

80 KIAS LAT

X.XX IPS AT YY.YY

80 KIAS VERT

X.XX IPS AT YY.YY

80 KIAS VERT

NOTE 1

NOTE 2

NOTE 3

NOTE 5: THE DISPLAY GOES BACK TO THE FIRST

REGIME WHEN THE ABOVE DATA HAS BEEN

COLLECTED

BMS ADVISORY LOG

BMS MAINTENANCE

NOTE 2

ACQUISITION

COMPLETE

NOTE 1

ACQUISITION

COMPLETE

NOTE 1

ACQUISITION

COMPLETE

NOTE 1

REDO 100% GND

PRESS REC

NOTE 5


FOR 4 SECONDS

NOTE 2

NOTE 3

Figure 7−29. Balance Monitor, Main Rotor Balance




BALANCE MONITOR COLLECT M/R

DATA RUN XX

MAIN ROTOR

BALANCE


100% GND LAT

X.XX IPS AT YY.YY

NOTAR�

BALANCE

BMS ADVISORY LOG

MAIN ROTOR TRACK

BMS FAULT LOG

BMS VERSION LOG

RUN XX M/R

MEASUREMENTS

MAIN ROTOR

CONFIGURATION

MAIN ROTOR

SOLUTION OPTIONS

DISPLAY M/R

SOLUTION RUN XX

HOVER IGE LAT

X.XX IPS AT YY.YY

80 KIAS LAT

X.XX IPS AT YY.YY

80 KIAS VERT

X.XX IPS AT YY.YY

120 KIAS LAT

X.XX IPS AT YY.YY

120 KIAS VERT

X.XX IPS AT YY.YY

NOT ACQUIRED

OR NOTE 1

NOTE 1: COULD APPLY FOR EACH REGIME

SPECTRUM

BMS MAINTENANCE


F92−085

Figure 7−30. Balance Monitor, Run M/R Measurements




BLADE 5 TRIM TAB

<XXX> MILS

NOTE 1: �ENT" KEY SELECTS DIGITS TO BE EDITED,

� AND � KEYS INCREASE/DECREASE DIGIT VALUE,

�REC" KEY STORES SELECTED VALUES, �CLR" EXITS

OUT OF MENU TO DISPLAY BLANK.


DATA RUN XX

MAIN ROTOR

BALANCE


NOTAR�

BALANCE

SPECTRUM

MAIN ROTOR TRACK RUN XX M/R

MEASUREMENTS

MAIN ROTOR

CONFIGURATION

MAIN ROTOR

SOLUTION OPTIONS

DISPLAY M/R

SOLUTION RUN XX

BLADE 1 HUB WT

<XXX> GRAMS

BLADE 5 HUB WT

<XXX> GRAMS

BLADE 1 PC WT

<XXX> GRAMS

BLADE 5 PC WT

<XXX> GRAMS

BLADE 1 TRIM TAB

<XXX> MILS

THROUGH

THROUGH

THROUGH

NOTE 1

NOTE 1 AND 2

NOTE 1 AND 2

NOTE 1 AND 2

NOTE 1

NOTE 1

NOTE 2: STEP THROUGH BLADES SEQUENTIALLY

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE


F92−086

Figure 7−31. Balance Monitor, Main Rotor Configuration





DATA RUN XX

MAIN ROTOR

BALANCE


NOTAR�

BALANCE

SPECTRUM


MEASUREMENTS

MAIN ROTOR

CONFIGURATION

MAIN ROTOR

SOLUTION OPTIONS

DISPLAY M/R

SOLUTION RUN XX

ADJUSTMENTS USED

<PCL/TAB/WEIGHT>

COMPUTE

<ENTIRE SOLTN>

NOTE 1

NOTE 2

NOTE 1

NOTE 2: OPERATOR OPTIONAL SELECTION

OR

OR

OR

OR

NOTE 2

NOTE 2

NOTE 2

<PCL/TAB>

<PCL WEIGHT>

COMPUTE

<GND SOLTN ONLY>

COMPUTE

<80 KIAS SOLUTION>


BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE

OR

<TAB/WEIGHT>

NOTE 1: �ENT" KEY SELECTS FIELD TO BE EDITED,

� AND � KEYS CHANGE FIELD SELECTION,

�REC" KEY STORES THE SELECTION, �CLR" EXITS


F92−087

Figure 7−32. Balance Monitor, Main Rotor Solution Options




TAB DWN XXX MILS

BLD X <NOT MADE>

TAB UP XXX MILS

BLD X <NOT MADE>

HUB ADD XXX.X G

BLD X <NOT MADE>

ENTIRE SOLTN

<NOT MADE>

80 KIAS SOLTN

<NOT MADE>

GND SOLTN ONLY

<NOT MADE>

HUB SUB XXX.X G

BLD X <NOT MADE>

NOTE 1: MESSAGE FLASHING IF COMPUTING A SOLUTION

NOTE 2: �ENT" KEY SELECTS FIELD TO BE EDITED, � AND � KEYS

CHANGE FIELD SELECTION FROM NOT MADE TO MADE, �REC" KEY STORES

SELECTION, �CLR" EXITS TO BLANK DISPLAY.


DATA RUN XX

MAIN ROTOR

BALANCE


NOTAR�

BALANCE

SPECTRUM


MEASUREMENTS

MAIN ROTOR

CONFIGURATION

MAIN ROTOR

SOLUTION OPTIONS


DISPLAY M/R

SOLUTION RUN XX

NOTE 3: SELECTIONS ARE NOT MADE, ALL MADE, OR AS SELECTED

PCL UP XX.X FLAT

BLD X <NOT MADE>

NOTE 3

PREDICTED VIBS

BELOW X.XX IPS

NOTE 2PCSE ADD XXX.X G

BLD X PCSE <NOT MADE>

OR

OR

OR

PCL DN XX.X FLAT

BLD X <NOT MADE>

OR

OR

OR

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE

COMPUTING M/R

SOLTN RUN XX.XXNOTE 1

AUTOMATIC

NOTE 4: SELECTIONS ARE ALL MADE, OR AS SELECTED

PCSE SUB XXX.X G

BLD X PCSE <NOT MADE>

NOTE 4

NOTE 4

NOTE 2

NOTE 2

NOTE 2

NOTE 2

NOTE 2

NOTE 2

NOTE 2

F92−088

Figure 7−33. Balance Monitor, Display M/R Solution




BALANCE MONITOR

FLASH STROBE

BLD SPREAD <ON>

MAIN ROTOR

BALANCE


NOTAR�

BALANCE

SPECTRUM

MAIN ROTOR TRACK

DEFAULT

OR

FLASH STROBE

BLD SPREAD <OFF>


BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE

F92−089

Figure 7−34. Balance Monitor, M/R Track




NOTE 1

NOTE 2

NOTE 3


BALANCE MONITOR COLLECT NOTAR

RUN XX RPM XXXX

MAIN ROTOR

BALANCE

NOTAR

BALANCE

SPECTRUM

MAIN ROTOR TRACK RUN XX NOTAR

MEASUREMENTS

NOTAR WEIGHT

CONFIGURATION

DISPLAY NOTAR

SOLUTION RUN XX

FLY 100% GND

PRESS REC

ACQUIRING

100% GND RADIAL

100% GND RADIAL

X.XX IPS AT YY.YY

ACQUISITION

COMPLETE

NOTE 2: WHEN COMPLETE, RESULT DISPLAYED FOR 4 SECONDS.

NOTE 1:AUTOMATICALLY STEPS THROUGH ACQUIRING


NOTE 3: WHEN COMPLETE, MESSAGE DISPLAYED FOR 1 SECOND

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE


F92−090

Figure 7−35. Balance Monitor, NOTAR� Balance




NOTE 3

NOTE 4

DISPLAY NOTAR

SOLUTION


BALANCE MONITOR

COLLECT

NOTAR DATA

MAIN ROTOR

BALANCE

NOTAR

BALANCE

MAIN ROTOR

TRACK

NOTAR

MEASUREMENTS

NOTAR WEIGHT

CONFIGURATION

100% GND RADIAL

X.XX IPS AT YY:YY

STD 1 WEIGHT

<XX.X> GRAMS

NOTE 1

PREDICTED VIBS

BELOW X.XX IPS

ADD XXX GRAMS

STD XX <NOT MADE>

NOTE 3

STD 13 WEIGHT

<XX.X> GRAMS

AUTOMATIC

NOT ACQUIRED

OR

OR

NOTAR SOLUTION

<NOT MADE>


COMPUTING NOTAR

SOLTN RUN XX.XX

SUB XXX GRAMS

STD XX <NOT MADE>

NOTE 1: �ENT" KEY SELECTS DIGITS TO BE EDITED,

� AND � KEYS INCREASE/DECREASE DIGIT VALUE,

�REC" KEY STORES SELECTED VALUES, �CLR" EXITS


NOTE 2: MESSAGE FLASHING IF COMPUTING A SOLUTION

NOTE 3: �ENT" KEY SELECTS FIELD TO BE EDITED, � AND � KEYS

CHANGE FIELD SELECTION FROM NOT MADE TO MADE, �REC" KEY STORES


NOTE 4: SELECTIONS ARE MADE, ALL MADE OR AS SELECTED

NOTE 1

NOTE 2

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE

SPECTRUM

F92−091

Figure 7−36. Balance Monitor, NOTAR� Data




BALANCE MONITOR MAIN ROTOR

BALANCE


FLY 100% GND

PRESS REC

FLY HOVER IGE

PRESS REC

FLY 120 KIAS

PRESS REC

NOTAR

BALANCE

SPECTRUM

MAIN ROTOR TRACK

BMS FAULT LOG

BMS VERSION LOG

TREND SPECTRUMS

RANDOM SPECTRUM

PRESS REC

GATHERING RANDOM

NO XX SPC XX/XX

RANDOM SPECTRUM

NO XX COMPLETE

NOTE 1

GATHERING TREND

SET XX SPC XX/XX

NOTE 1

TREND SPECTRUM

SET XX COMPLETE

NOTE 2

NOTE 1

NOTE 2

NOTE 1

NOTE 2

NOTE 1: AUTOMATICALLY STEPS THROUGH ACQUIRING



FOR 3 SECONDS AND DISPLAY GOES BACK TO

RANDOM SPECTRUM MENU


FOR 3 SECONDS AND THE DISPLAY GOES TO THE

NEXT REGIME.

GATHERING TREND

SET XX SPC XX/XX

TREND SPECTRUM

SET XX COMPLETE

GATHERING TREND

SET XX SPC XX/XX

TREND SPECTRUM

SET XX COMPLETE

NOTE 2


F92−092

Figure 7−37. Balance Monitor, Spectrum





BALANCE


NOTAR

BALANCE

SPECTRUM

MAIN ROTOR TRACK

BMS OK

DBASE USAGE XXX%

ADVISORIES = XX

END OF BMS

FAULT LOG

OR

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE

BALANCE OK

END OF BMS

ADVISORY LOG

OR

NOTE 1: OR VIBRATION DATA

NOTE 2: OR ERROR MESSAGES

NOTE 1

NOTE 2


F92−093

Figure 7−38. Balance Monitor, BMS Fault Log





BALANCE


BMS BOARD P/N

XXXXX−XX

BMSBP VER XX.XXX

P/N XXXXX−XX

BMSBP CHECKSUM

XXXXXXXX

NOTAR

BALANCE

SPECTRUM

MAIN ROTOR TRACK

MAIN ROTOR MODEL

VER XX.XXX

BMSOP VER XX.XXX

P/N XXXXX−XX

BMSOP CHECKSUM

XXXXXXXX

NOTAR MODEL

VER XX.XXX

VIB MONITOR

VER XX.XXX

SPECTRUM SETUP

VER XX.XXX

BMSBM CHECKSUM

XXXXXXXX

BMSBM VER XX.XXX

P/N XXXXX−XX

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE


F92−094

Figure 7−39. Balance Monitor, BMS Version Log





BALANCE


CLEAR FAULT

LOG

CLEAR ADVISORY

LOG

CLEAR SPECTRUM

LOG

NOTAR

BALANCE

SPECTRUM

MAIN ROTOR TRACK

FORMAT DATABASE

AND RESET BMS

CLEAR M/R

BALANCE LOG

CLEAR NOTAR

BALANCE LOG

CLEAR ALL LOGS

CLEAR SETUP

CONFIGURATION

BMS ADVISORY LOG

BMS FAULT LOG

BMS VERSION LOG

BMS MAINTENANCE


NOTE 1: FOR ALL ‘‘CLEAR’’ MENU SELECTIONS, PRESS ENT KEY AND A ‘‘PRESS TO

CLEAR’’ MESSAGE WILL BE DISPLAYED. PRESS REC KEY TO CLEAR THE LOG

AND A ‘‘CLEARED OK PRESS ANY KEY’’ MESSAGE WILL BE DISPLAYED.

PRESSING ANY KEY WILL RETURN TO THE ‘‘CLEAR LOG MENU’’.

NOTE 2: PRESS ENT KEY AND A ‘‘PRESS TO FORMAT AND RESET’’ MESSAGE WILL

BE DISPLAYED. PRESS REC KEY TO FORMAT THE DATA BASE AND A ‘‘DBASE

FORMATTED INITIALIZING BMS’’ MESSAGE WILL BE DISPLAYED. PRESSING

ANY KEY WILL RETURN TO THE ‘‘CLEAR LOG’’ MENU AFTER 30 SECONDS

NOTE 1

NOTE 2

F92−095

Figure 7−40. Balance Monitor, BMS Maintenance




AIRCRAFT MONITOR EXCEEDANCE LOGENT ENT

EXCEED LOG XXX

L ENG TORQUE

DATE XX−XX−91

TIME XX:XX:XX

PEAK VALUE XXX%

SEC TO PK = XX SEC

DATA LOG NO X

SEC ABV T1 = XX

SEC TO T2 = XX

SEC ABV T2 = XX

SEC TO T3 = XX

SEC ABVT3 = XX

SEC TO T4 = XX

SEC ABV T4 = XX

NOTE: PRESSING THE BUTTON WILL TAKE YOU TO THE PREVIOUS HIGHEST MENU LEVEL.MENU

F92−096


TREND LOG

FAULT LOG

IIDS SETUP

THE EXCEEDANCE LOG PROVIDES A ‘‘SNAPSHOT" RECORD OF THE PARAMETER DATA AT APARTICULAR MOMENT IN TIME. THIS TYPE OF RECORD OCCURS WHENEVER A PARAMETEREXCEEDANCE IS DETECTED.

Figure 7−41. Aircraft Monitor, Exceedance Log Menu




AIRCRAFT MONITORENT

ENTTREND LOG XX

LEFT ENGINE

DATE

TIME

NP XXX%

T1 XC

TORQUE XX%

NG XX%

EGT XXXC

P0 XXXX FT

OAT XX C

PERFORM MARGIN

L PA CK NG −XX

L PA CK EGT−XX

NG COR FTR XX.X

EGT COR FTR XXX

NOTE: PRESSING THE BUTTON WILL TAKE YOU TO THE PREVIOUS HIGHEST MENU LEVEL.MENU

F92−097

EXCEEDANCE LOG

TREND LOG

FAULT LOG

IIDS SETUP

Figure 7−42. Aircraft Monitor − Trend Log




AIRCRAFT MONITORENT

ENTFAULT LOG

L ENG S/N

R ENG S/N

DATE

TIME

IIDS FALT1=X XX

XX, XX

IIDS FALT XX

ACFT FALT=X

SENS FALT=X

BMS FALT=X

LEFT EEC

DSCWD1 =X XX XX

NCFUR1=X XX

NCFUR2=X XX XX

NCFUR3=X XX

CFUR=0

RIGHT EEC

DSCWD1=

RIGHT EEC MENU SAME AS

LEFT EEC MENU.

NOTE: PRESSING THE BUTTON WILL TAKE YOU

TO THE PREVIOUS HIGHEST MENU LEVEL.

MENU


F92−098

EXCEEDANCE LOG

TREND LOG

FAULT LOG

IIDS SETUP

THE FAULT LOG CONTAINS DATA ASSOCIATED WITH EECFAULTS AND FAILURES DETECTED IN THE IIDS, BMS, ORAIRCRAFT TRANSDUCERS/SENSORS. THIS LOG IS RE-CORDED WHENEVER AN IIDS OR EEC FAULT IS DE-TECTED. STORAGE IS AVAILABLE FOR 100 FAULT LOGS.

Figure 7−43. Aircraft Monitor, Fault Log Menu





AIRCRAFT MONITOR EXCEEDANCE LOG

ENG INSTALL PWC

IPS INSTALLED

HT/DEFOG INSTAL

ROTOR BRK INSTAL

FAULT LOG

IIDS SETUP

TREND LOG

FWD FUEL CAL XXX

AFT FUEL CAL XXX

TOP LVL SFTWR PN

XXXXXXXXXXXX

A/C NO XXXXXXXX

CFG DAT MM−DD−YY

CFG TIM HR:MN:SE

OPER SFTWR PN

XXXXXXXXXXXX

MAINT SFTWR PN

XXXXXXXXXXXX

BMSOP SFTWR PN

XXXXXXXXXXXX

BMSBP SFTWR PN

XXXXXXXXXXXX

SETUP DATA ID

XXXXXXXX

BMSBM SFTWR PN

XXXXXXXXXXXX


F92−099

Figure 7−44. Aircraft Monitor − IIDS Setup




FUEL CALIBRATION FWD FUEL CAL<XXX>

AFT FUEL CAL <XXX>

SET CAL CODE


DO CALIBRATION AIRCRAFT READY?

CRUISE ATTITUDE?

NOTE 1: PRESS �ENT" FOR MORE THAN 4 SECONDS TO ENTER

FUNCTION, THIRD LEVEL MENU APPEARS.

NOTE 2: �ENT" SELECTS DIGITS TO BE EDITED, AND KEYS

CHANGE SELECTED DIGITS, �REC" KEY STORES


NOTE 3: PRESS �ENT" FOR MORE THAN 4 SECONDS COMMANDS

CALIBRATION. IIDS WITH CAL CODES AFTER

CALIBRATION COMPLETE. �REC" CHANGES CODE TO

CALCULATED VALUE. PRESS ‘‘MENU’’ TWICE TO

RETURN TO TOP LEVEL

FWD FUEL CAL XXX

AFT FUEL CAL XXX

NOTE 1

NOTE 1

NOTE 2

NOTE 3

F92−100


NOTE: TO MOVE HORIZONTALLY ( → ) TO THE NEXT LOWER LEVEL − PRESS ENT

Figure 7−45. Fuel Calibration




SET ENGINE PARM LNG COR FCTR <XX.X>

LEGT CORFCT <XX.X>


RNG COR FCTR <XX.X>

REGT CORFCT <XX.X>

NOTE 1

NOTE 1

NOTE 1: �ENT" SELECTS DIGITS TO BE EDITED, � AND � KEYS

CHANGE SELECTED DIGITS, �REC" KEY STORES


F92−101

TOP LEVEL SECOND LEVEL

Figure 7−46. Set Engine Parameters

SET TIME/DATE


NOTE 1

F92−102


TIME <HH:MM>

DATE MM−DD−YYNOTE 1.: ‘‘ENT’’ KEY SELECTS FIELD TO BE SET (MINUTE,

HOURS, DAY, MONTH, YEAR) AND SELECTED FIELDBLINKS, � AND � KEYS INCREMENT/DEINCRE-MENT DIGIT VALUE, ‘‘REC’’ KEY CHANGES TIMEAND DATE TO SELECTED VALUES

Figure 7−47. Set Time/Date

Handling Servicingand Maintenance



S E C T I O N VIIIHANDLING, SERVICING

AND MAINTENANCETABLE OF CONTENTS

PARAGRAPH PAGE8−1. Hoisting, Lifting, and Jacking 8−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−2. Towing and Moving 8−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−1. Helicopter Towing and Ground Handling 8−4. . . . . . . . . . . . . . . . . . . . .

8−3. Parking and Storage 8−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−2. Helicopter Tiedowns and Covers 8−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−3. Helicopter Grounding 8−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−4. Access and Inspection Provisions 8−10. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−4. Access Methods 8−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−5. Nose Access Panels 8−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−6. Left Side Access Panels 8−13. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−7. Right Side Access Panels 8−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−8. Top View Access Panel 8−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−9. Bottom View Access Panels 8−16. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−10. Stabilizers Access Panels 8−17. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−11. Cabin Floor Interior Access Panels 8−18. . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−12. Pedestal Access Panels 8−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−13. Fan Assembly Access Panels 8−20. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−5. Servicing 8−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−14. Servicing Points 8−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 8−1. Acceptable Fuels 8−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 8−2. Servicing Materials 8−22. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Figure 8−15. Fuel System Gravity Filler Port 8−25. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−16. Hydraulic System 8−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−17. Main Transmission Servicing 8−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−18. Rotor Brake 8−30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−19. Engine Oil System − Servicing 8−32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−6. Aircraft Cleaning 8−33. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−7. Cockpit Door Removal 8−34. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Handling Servicingand Maintenance


Original 8−iiReissue 1

PARAGRAPH PAGEFigure 8−20. Cockpit Door Attachment 8−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−8. Cabin Seats: Removal/Installation 8−36. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−21. Cabin Passenger Seat Attachment 8−36. . . . . . . . . . . . . . . . . . . . . . . . . .

8−9. Copilot Flight controls 8−37. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−22. Copilot Pedals 8−37. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

8−10. Engine Charts 8−38. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−23. EGT Vs Time − All Conditions Except Starting 8−38. . . . . . . . . . . . . . .

Figure 8−24. EGT Vs Time − Starting 8−39. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−25. Power Turbine (NP) Speed Vs Time 8−39. . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−26. Compressor Turbine (NG) Speed Vs Time 8−40. . . . . . . . . . . . . . . . . . . .

Figure 8−27. Engine Overtorque Limits − All Conditions 8−40. . . . . . . . . . . . . . . . . .

8−11. Special Operational Checks and Procedures 8−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Engine NP Overspeed Test Procedure 8−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Engine Out/Low Rotor Warning Check 8−41. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Hydraulic System Check 8−42. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

VSCS Check 8−42. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Wet Engine Motoring Run 8−43. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dry Engine Motoring Run 8−43. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Engine Wash procedures 8−44. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−28. Engine Wash Panel 8−46. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Manual Engine Shutdown Check 8−47. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Manual Engine Start Check 8−48. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Autorotation RPM Check 8−49. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Battery Removal 8−49. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Resetting IIDS Time/Date 8−50. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 8−29. Set Time/Date 8−50. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


MD900 (902 Configuration with PW 207E) Handling Servicingand Maintenance


SECTION VIIIHANDLING, SERVICING, AND

MAINTENANCE

8−1. HOISTING, LIFTING, AND JACKING

Hoisting, lifting, and jacking of the helicopter shall only beperformed with the proper equipment and tools as specified in theMD900 Rotorcraft Maintenance Manual. Failure to follow thespecified procedures will result in damage to aircraft components.

8−2. TOWING AND MOVING

Moving the helicopter on prepared surfaces is accomplished by mounting groundhandling wheels to fittings located on the landing gear skid tubes.

The ground handling wheel set is used for moving the MD Explorer by hand andfor towing. The wheels are manually lowered with a detachable jack handle, andare held in the down position by a mechanical lock. The ground handling wheelset is equipped with a tow bar attach fitting.

Helicopter Manual Moving:

Ensure all stress panels listed in Figure 8−11 are installed.

Airframe structure damage can occur if stress panels are not inplace before moving helicopter.

NOTE: The wheel set attaches at four points, two inboard and two outboard, on the skidassemblies.

A ‘‘T" handle is strapped to the skid tubes and extends out, to pull the wheelsto and from the helicopter.

Position wheel set over skid tubes and roll wheel set forward.

NOTE: The wheel set can be installed in either direction, depending on jack handleposition.

Attach wheel set to attach points on skid tubes.

Hold tail up while lowering ground handling wheels.

CAUTION

CAUTION


MD900 (902 Configuration with PW 207E)Handling Servicingand Maintenance


Jack hydraulic ram which forces wheels down and skids up.

When the jack is extended, a mechanical safety latch automatically snaps intoposition, to prevent the wheels from going up in the event of loss of hydraulicpressure.

Operators and maintenance personnel should avoid lead−lag loadsin excess of 25 lb (11 kg) at the tip of the main rotor blades.

Excessive lead−lag load applied to the main rotor blades during ground handlingcan result in damage to the damper assembly.

Manually move helicopter on ground handling wheel set by balancing tailboomand pushing on rear fuselage portion of airframe.

When ground handling helicopter do not lift main rotor blades toclear objects. The main rotor should be rotated to clear objects.To prevent rotor component damage, the main rotor hub deflectionfor a non−operating rotor is not to exceed four feet up, maximum.Measurement to be taken from static rest.

Helicopter Towing

The towbar is equipped with caster wheels and is designed for use with theground handling wheels and allows the helicopter to be moved by one person.

The towbar does not interfere with equipment that may be hung under the heli-copter



Raise helicopter up with wheel set.

Position caster wheels, to straddle, over front skid tubes.

Attach nylon strap under skid tubes and ratchet skid tubes into rubber cups.

Attach tow bar to a power unit.

Do not tow helicopter at speeds over 5 mph.When ground handling helicopter do not lift main rotor blades toclear objects. The main rotor should be rotated to clear objects.To prevent rotor component damage, the main rotor hub deflectionfor a non−operating rotor is not to exceed four feet up, maximum.Measurement to be taken from static rest.

CAUTION

CAUTION

CAUTION

CAUTION




Avoid sudden stops and starts.

Avoid short turns, which could cause helicopter to turn over.

A safe minimum turning radius is approximately 20 feet (6.1m).

Allow inside wheel to turn (not pivot) while helicopter is being turned.

Helicopter Transport

The heli−porter is designed for the MD Explorer with the use of a towing tractoror vehicle.

The heli−porter is a welded tubular steel frame with dual pneumatic swivelcaster on the front and rear. The platform is a grated walkway with hold downsafety straps for the landing gear. The heli−porter has a hand brake to the reartires and has a 10,000 lb (4540 Kg) capacity.



Land or hoist helicopter (Ref. RMM, Section 07−10−00) on heli−porter platform.

Attach safety hold−down straps to skid tubes.

Release heli−porter hand brake.

When ground handling helicopter do not lift main rotor blades toclear objects. The main rotor should be rotated to clear objects.To prevent rotor component damage, the main rotor hub deflectionfor a non−operating rotor is not to exceed four feet up, maximum.Measurement to be taken from static rest.

Do not tow helicopter at speeds over 5 mph. A safe minimum turningradius is approximately 20 feet (6.1m).

Attach heli−porter hook−up to a tow vehicle.

CAUTION

CAUTION

CAUTION

CAUTION




VIEW ROTATED

F92−103

SKID TUBE

GROUNDHANDLING

ATTACH POINTS

HELI−PORTER

TOW BAR

Figure 8−1. Helicopter Towing and Ground Handling

8−3. PARKING AND STORAGE

Helicopter tiedowns and covers

Covers and tiedowns (Ref. Figure 8−2) are provided to shield the MD Explorerfrom inclement weather conditions and other outside environmental factors thatcould cause FOD damage while the helicopter is parked, moored, or while instorage.

NOTE: The decision to use protective covers and tiedowns is determined by theprevailing weather conditions, length of storage/parking, and location.

Forward and aft tiedowns

Each tiedown (Ref. Figure 8−2) has a quick connect fitting with a streamerattached ‘‘REMOVE BEFORE FLIGHT". Two aft tiedown straps are tobe attached to the upper aft fitting. Two separate upper forward tiedownsattach to the forward fuselage. Additional lower fore and aft tiedown attachpoints are located on left and right side of helicopter.

Blade tiedowns

Blade tiedowns (Ref. Figure 8−2) are socks, which fit over the blade tip, withor without the blade covers installed.

Each blade tiedown is fitted with a generous length of rope which can betied down at any convenient spot.

Upper deck cover

The upper deck cover (Ref. Figure 8−2), encloses the NOTAR� inlet, particleseparator inlets and exhaust stacks.




Attach cover, at forward corners to snap−head screws placed in existing screwholes on lower edge of particle separator inlet.

Attach cover, at rear corners with similar snaps, or with a strap going under-neath the tailboom where it meets the fuselage.

NOTAR� boom cover

The boom cover (Ref. Figure 8−2) is a tubular cover made of nylon. Attachboom cover to thruster using side−release buckles.

NOTAR� thruster cover

The MD Explorer thruster cover (Ref. Figure 8−2) encloses the thruster coneand chokes tightly around the base near the trailing edge of the horizontalstabilizor.

Position cover on thruster.

Tighten cover with strap assembly.

Pitot tube cover

The pitot cover (Ref. Figure 8−2) is vinyl and reinforced with galvanized steelstaples at stress points.

A bright red warning streamer, ‘‘REMOVE BEFORE FLIGHT" attachesto the bottom edge of the cover

Attach pitot cover around pitot base.

Bubble cover

The MD Explorer bubble cover (Ref. Figure 8−2), encloses the entire canopy,including the windshield, front and rear doors and windows.

The cover, attaches at four points.

The cover, is color−coded, with swatches sewn in the corners, for ease of instal-lation. Red = Left, Green = Right.

Attach upper rear corners to snap−head screws placed in existing screw holeson lower edge of particle separator inlet.

Attach straps at lower rear corners to rear struts.

Tighten special rope in top and bottom hems, to insure a guarantee againstwind chaffing.

A large bright red pocket is sewn in the cover, for the temperature probe.




UPPER DECK COVER

F92−104ABUBBLE COVER

BLADE COVER

BOOM COVER

ENGINE AREA COVER

THRUSTER COVERROTOR HUB COVER

AFT UPPER TIEDOWNFORWARD UPPER TIEDOWN

BLADE TIEDOWN

PITOT TUBE COVER

AFT LOWER TIEDOWN(IF INSTALLED)

FORWARD LOWER TIEDOWN(IF INSTALLED)

Figure 8−2. Helicopter Tiedowns and Covers




Rotor hub cover

The MD Explorer rotor hub cover (Ref. Figure 8−2) overlaps with the bladecovers and the engine area cover to insure complete protection for the entiremain rotor system.

Position cover over top of rotor hub.

Attach cover with buckles under each blade.

Tuck lower part of cover into the aperture beneath the rotor.

Engine area cover

The engine area cover is similar and does the same job as the upper deckcover, except that it also encloses the fan inlet.

Blade covers − standard

Blade covers (Ref. Figure 8−2) can be installed from the ground.

Tighten covers at blade root with attached straps and buckles.

The small opening at the blade tip bottom allows attachment of tiedown ropes.

Blade covers − cold weather

The MD Explorer cold weather blade covers are similar to the standard bladecovers, but are fitted with full length zippers and heater hose boots nearthe blade root.

Helicopter Parking

The decision to use protective covers and tiedowns is determined by the prevailingweather conditions, length of parking, and location.

Normal Conditions

NOTE: Tie down rotor blade(s) whenever helicopter is parked in an area subject toturbulent or gusting winds to prevent rotor windmilling. The maximum blade tiedown load is when the blade tip just begins to deflect downward.

Install pitot cover.

Close and secure all doors, windows and access panels.

Statically ground helicopter if possible.

Turbulent/Gusting Wind Conditions

Tie down all blades in winds of 45 knots or more to preventexcessive flapping and possible flexbeam damage.

NOTE: Maximum demonstrated wind speed for starting and stopping the rotor is 50knots.

CAUTION




If possible, park helicopter into the prevailing wind and secure in accordancewith normal parking conditions.

Statically ground helicopter (Ref. Figure 8−3).

F92−105

GROUNDING JACK(2 LOCATIONS)

CABLE PLUG

Figure 8−3. Helicopter GroundingHelicopter Storage

Install protective covers as necessary (Ref. Figure 8−2) to prevent entry of blowingdust, water, freezing rain, snow and/or foreign objects into the helicopter duringground storage.

Maintain full fuel cell to reduce condensation in the cell.

Ensure drain holes are free of debris and kept open during storage period.

Store helicopter in hangar, if space allows.

Statically ground helicopter (Ref. Figure 8−3).

Flyable Storage

Flyable storage will maintain a stored MD Explorer in an operable condition.If daily use is impractical, the following procedures will keep the helicopterin flyable condition. At regular intervals inspect helicopter. Date and typeof storage must be recorded in helicopter records.

Perform engine run−up, at least once every five days.

Perform pre−flight inspection, at least once every seven days.

Inspect helicopter and treat for corrosion control.

Inspect static ground wires, blade tiedowns and mooring devices at regularintervals.

Inspect tiedowns immediately after winds exceeding 35 knots.

Enter type of storage and date helicopter was placed in storage, in helicopterrecords.




High Wind Conditions − Helicopter Mooring

When severe storm conditions or wind velocities higher than 40 knots are forecast,helicopter should be hangared or evacuated to a safe area. If the helicopter mustbe parked in the open during high winds, comply with the following.

Structural damage can occur from flying objects during high windconditions. Helicopter should be hangared or evacuated to a safeweather area when wind conditions above 75 knots are expected.

If a paved ramp with tiedown rings are available, park helicopter headed indirection from which highest forecast winds are expected.

Secure helicopter to ramp tiedowns using forward and aft tiedowns (Ref.Figure 8−2).

If a paved ramp with tiedown rings are not available, park helicopter on anunpaved parking area, headed in the direction from which highest forecast windsare expected.

Install blade tiedowns.

Tiedown rotor blades, whenever helicopter is parked, to preventrotor damage from blade flapping as a result of air turbulence fromother aircraft or wind gusts. The maximum blade tie down loadis when the blade tip just begins to deflect downward.

Install engine area cover (Ref. Figure 8−2), and pitot cover.

Fill fuel cell, if possible.

After winds subside, inspect helicopter carefully for damage which may havebeen inflicted by flying objects.

Return to service

Flyable Storage Depreservation and Activation

Remove protective covers and tiedowns.

Clean helicopter as necessary.

Open all doors and ventilate helicopter.

Record date helicopter was prepared for service in helicopter records.

Remove static ground wire installed for storage.

Perform preflight checks.

CAUTION

CAUTION




8−4. ACCESS AND INSPECTION PROVISIONS

Various doors, covers, panels, and fairings are located through out the airframeto provide access for inspection, maintenance, and servicing. External and internaldoors, covers, panels and fairings are shown in Figure 8−5 thru Figure 8−13. Eachdoor, cover, panel, and fairing has a letter and a number designator. Each figureis directly related to a corresponding table which lists a reference designator, panelname, accessible item, access method and fastener type.

Reference Designator:

The number indicates the nearest attaching fuselage station.

The letter indicates the location:

(N) Nose(L) Left Side(R) Right Side(T) Top Side(B) Bottom Side(A) Cabin Floor (interior access)(S) Stabilizer(P) Pedestal (interior)(F) Fan Assembly (interior)

A combination of two letters may be used to help identify a door, cover, panelor fairing:

(FR) floor right(FL) floor left

L and R will indicate doors, panels, and covers at the same station location:

F(L/R)160.0.

Removal and Installation Methods:

Removal or installation of doors, covers, panels, and fairings are described bya method listed in a table with a supporting illustration. The type of fastenerand quantity used to remove or secure the door, cover, panel, and fairing is listed.




9G06−008F92−106

UNLOCK−COUNTERCLOCKWISELOCK−CLOCKWISE

LOCKED

PUSH HERETO UNLOCKROTATES 180°

TO UNLOCK

1/4 TURN

UNLOCK− COUNTERCLOCKWISELOCK−CLOCKWISE

UNLOCKED

1/4 TURN

SCREW

BOLT

PULL TO RELEASEDOOR PINS

SLEEVE BOLT

LEVER ACTION HANDLE

CAMLOC

2 PULL

1 PUSH

KEYLOC CAMLOC LATCH

UNLOCKED

HINGE

LOCKED UNLOCKED

PUSH TOUNLOCK PUSH TO

LOCK

LATCHCAMLOC

LOCKED

HINGE PIN PIN CAMLOC

HOOK−1

LIFT

LOCK−TWO ACTIONS

PULL

UNLOCK−ONE ACTION

PUSH−2

LATCH

PULL OFFHERE

STRUT

LIFTHERE

TURN AND LOCK

CLOSED UNLOCKEDLOCKED

Figure 8−4. Access Methods




9G06−001F92−107

N80

N82

N106

ItemNo. Name Permits Access To

Removal and Installation

Quantity TypeMethod

Ref.Figure 8−4

N80 Nose Door Pitot − Static System, Battery 21

LatchKeyloc

BA

N82 Panel Landing Light, Flight Control RodsAvionics Cooling Fan, Wire Harness,External Power Box

18 Screw L

N106 Panel 30 Screw L

Figure 8−5. Nose Access Panels




F927−092

L155 L210

L220 L260

L270

L262

L166

L107L109 L240



Quantity TypeMethod

RefFigure 8−4

L107 Crew Door Copilot Instrument Panel, Pedestal Console 111

Lever Action HandleHinge Pin / PinStrut End

HJ, K

T

L109 Access Panel AssemblyAvionics, LH

Electrical, Avionics, Flight Controls,Static System Drain Valve

114

KeylocCamloc

AB

L155 Forward Access DoorAssembly, LH

Main Transmission Access, FlightControl Actuators, Hydraulic HandPump, System 1 Hydraulic Manifold/Reservoir

52

CamlocHinge

E D

L166 Passenger Door, LH Passenger and Cargo Compartment 1 Lever Action Handle HL210 Transmission Access

Door Assembly, LHUPPER W/ NACAINLET

Main Transmission Access, EngineReduction Gearbox Housing

25

HingeCamloc

DE

L210 Transmission AccessDoor Assembly, LHLOWER W/ NACAINLET

Main Transmission Access, EngineReduction Gearbox Housing

7 Camloc E

L220 Engine Air Inlet PanelAssembly, LH

Inlet, Engine Compressor 25 Fastener Sleeve Bolt G

L240 GPU/EPR Door (AFT) Auxiliary / External Power Recep-tacle

1 Camloc E

260 Engine Cowling Assem-bly, LH

Engine, Engine Controls 10 Fastener Sleeve Bolt G

L262 Baggage CompartmentDoor

Baggage Compartment,behind trim panels; Engine EEC’s,Wire Harness, Drain Lines, ElectricalLoad Center, Engine Fire Extinguish-ing Bottles (optional)

1 Lever Action Handle H

L270 Exhaust Ejector CowlAssembly, LH

Engine Exhaust 18 Machine Screw L

Figure 8−6. Left Side Access Panels




9G06−003F92−109

R270

R260 R220 R210

R155

R107

R109R128R158R185

R166

ItemNo.

Name Permits Access To


Quantity TypeMethod

RefFigure 8−4

R107 Crew Door Pilot Instrument Panel, Pedestal Consol 111

Lever Action HandlePin AssemblyStrut

HJ, K

T

R109 APU/EPR Door Auxiliary / External Power Recep-tacle

1 Camloc E

R128 Avionics Access Panel Electrical, Avionics, Flight Controls,ECS Bleed Air Lines, Static SystemDrain Valve

114

Key LocCamloc

AB

R155 Forward Access DoorAssembly, RH

Main Transmission Access, FlightControl Actuators, System 2Hydraulic Manifold/Reservoir

52

CamlocHinge

ED

R158 Fuel Cap and Adapter Fuel Filler Neck 1 Turn and Lock SR166 Passenger Door, RH Passenger and Cargo Compart-

ment11

Lever Action HandleRelease Pin

HR

R185 Fuel Drain Access PanelAssembly

Fuel Sump Drain Control, Cables 1 Camloc N

R210 Transmission AccessDoor Assembly, RHUPPER W/ NACA INLET

Main Transmission Access,Engine Reduction GearboxHousing

25

HingeCamloc

DE

R210 Transmission AccessDoor Assembly, RHLOWER W/ NACAINLET

Main Transmission Access,Engine Reduction GearboxHousing

7 Camloc E

R220 Engine Air Inlet PanelAssembly, RH

Inlet, Engine Compressor Section 25 Sleeve Bolt G

R260 Engine Cowl Assembly, RH

Engine, Engine Controls 11 Sleeve Bolt G

R270 Exhaust Ejector CowlAssembly, RH

Engine Exhaust 19 Machine Screw L

Figure 8−7. Right Side Access Panels




T155 T220

T240

9G06−009F92−110

TR218T292

TL218



Quantity TypeMethod

RefFigure 8−4

T155 Forward Fairing As-sembly Swashplate

Hydraulic Servo Actuators,Swashplate, Mixer

194

ScrewBolt

LM

TL118 Oil Dipstick HandHold, LH

Engine Oil Level and Filler 12

HingeCamlock

QE

TR118 Oil Dipstick HandHold, RH

Engine Oil Level and Filler 12

HingeCamlock

QE

T220 Aft Fairing AssemblySwashplate

Mast Support, Transmission,Flight Controls, ECS, EngineOil Level and Filler

584

ScrewBolt

LM

T240 Upper Inlet Duct As-sembly

Fan driveshaft, air inlet to fan 1924

ScrewBolt

LM

T292 Upper Tailboom Fair-ing Assembly

Required Panel Removal T240,L270 and R270

7 Screw L

Figure 8−8. Top View Access Panel




9G06−005F92−111

B142B178

B230



Quantity Type Method RefFigure 8−4

B142 Access Panel Assem-bly Center

Throttle Interconnect Cable, RHCollective

14 Screw L

B178 Access Panel Assem-bly Sump

Fuel Sump Drain Valves 22 Screw L

B230 Aft Crosstube CoverAssembly

Landing Gear Crosstube Aft 30 Screw L

Figure 8−9. Bottom View Access Panels




9G06−004F92−112

SR5SL5

S4

S6

SR3

SR1

SL3S2SL1

SR7

SL7



Quantity TypeMethod

RefFigure 8−4

S2 Lower Tailboom/Thruster Fairing Assembly

Horizontal Stabilizer Mount FittingStationary Thruster Mounting

22 Screw L

S4 Thruster End Cover Attachment Bolts For RotatingThruster

8 Screw L

S6 Leading Edge CoverCenter

10 Screw L

SL1 Upper Tailboom/Thruster Fairing As-sembly, Left Side


21 Screw L

SL3 Outboard Fairing As-sembly, LH (Endplate)

Vertical Stab Torque Tube, ControlRod Electrical Wiring, Position Light

14 Screw L

SL5 Center Access Cover,LH (Horizontal Stabi-lizer)

Wiring 10 Screw L

SR1 Upper Tailboom/Thruster Fairing Assembly Right Side


21 Screw L

SR3 Outboard Fairing As-sembly, RH (Endplate)

Vertical Stab Torque Tube, ControlRod Electrical Wiring, Position Light

14 Screw L

SR5 Center Access Cover(Horizontal Stabilizer)

Wiring 10 Screw L

SR7 Access Cover RH VSCS Actuator 10 Screw L

SL7 Access Cover LH VSCS Actuator 10 Screw L

Figure 8−10. Stabilizers Access Panels




AR129

AR138 AR155 AR165 AR230 AR250

AL250

A235

AL230

A217AL165 A170A160AL155

AL138

AL129

9G06−010F92−113

NOTE: ALL PANELS ARE STRESS PANELS.



Quantity Type Method RefFigure 8−4

A160 Cabin Floor ForwardPanel Assembly

Fuel Cell 69 Screw L

A170 Cabin Floor CenterPanel Assembly

Fuel Cell 89 Screw L

A217 Cabin Floor Aft PanelAssembly

Fuel Cell Aft Vent Rollover Valves 75 Screw L

A235 Baggage Floor CenterPanel

Condenser Fans, Condenser 64 Screw L

AL129 Cockpit Outboard LeftFloor Panel

LH Collective Stick Socket, Wire Harness, LHStatic Port

301

ScrewBolt

LM

AL138 Cockpit Floor LeftAccess Panel

Flight Control Tubes, Cyclic Bellcrank 241

Screw Bolt

LM

AL155 Cabin Floor Left For-ward Access Cover

Left Forward Fuel Vent Valve, Fuel Cell 12 Screw L

AL165 Cabin Floor Left Out-board Panel

Fuel Cell Frangible Valve, Wire Harness 76 Screw L

AL230 Cabin Floor Left Aft Ac-cess Cover

Left Aft Fuel Vent Valve, Fuel Cell 13 Screw L

AL250 Baggage Floor LeftOutboard Panel

Fuel Tee Fittings, Fuel Pressure Switch DrainTubing, Fuel Hose Shrouds, Fuel Catch Can

254

ScrewBolt

LM

AR129 Cockpit Outboard RightFloor Panel

Fire Overheat Bleed Air Leak Control, WireHarness, RH Static Port

341

ScrewBolt

LM

AR138 Cockpit Floor RightAccess Panel

Flight Control Tubes, Bellcranks, ThrottleCables

35 Screw L

AR155 Cabin Floor Right For-ward Access Cover

Right Forward Fuel Vent Valve, Fuel Cell 12 Screw L

AR165 Cabin Floor Right Out-board Panel

Fuel Cell Frangible Valve, Heat/Defog BleedAir Line, Fire Overheat Bleed Air Leak Detec-tor

76 Screw L

AR230 Cabin Floor Right AftAccess Cover

Right Aft Fuel Vent Valve, Fuel Cell 13 Screw L

AR250 Baggage Floor RightOutboard Panel

Fuel Hose Shrouds, ECS Tubing, StrobePower Supply,

254

ScrewBolt

LM

Figure 8−11. Cabin Floor Interior Access Panels




9G06−012F92−114

PR120

PL120



Quantity TypeMethod

RefFigure 8−4

PL120 Panel Wire Harness, Forward InterconnectPanel (Relays), Ground Modules

9 Camloc U

PR120 Panel Wire Harness, Forward InterconnectPanel (TB2, TB3,TB4)

9 Camloc U

Figure 8−12. Pedestal Access Panels




REF. STATOR

9G06−013F92−115

REF. FAN SUPPORT ANDHOUSING

F1 F2

F3

F4F5

F6

F7F9

F8



Quantity TypeMethod

RefFigure 8−4

F1 Anti−Torque Drive ShaftCover

Fan Drive Shaft 6 Screw L

F2 Anti−Torque Lower InletDuct Assembly

Plenum Fan Assembly 1914

ScrewBolt

LM

F3 Anti−Torque Fan Fair-ing/Center Body Assem-bly

Fan Assembly, Fan Driveshaft Cou-pling, Support Housing, Fan BalanceMonitor System Magnetic Pickup andAccelerometer

19 Screw L

F4 Anti−Torque Middle InletDuct Assembly

Fan Assembly Plenum air Inlet 419

ScrewBolt

LM

F5 Anti−Torque Fan UpperDuct Assembly

Plenum Air Inlet, Upper Stator Bladesattached

24 Screw L

F6 Upper Center Diffuser Upper Stator Blades attached 24 Screw L

F7 Lower Center Diffuser Fan Assembly and Diverter, LowerStator Blades attached

20 Screw L

F8 Lower Access Panel As-sembly

Diverter 4 Latch P

F9 Anti−Torque Fan LowerDuct Assembly

Fan Assembly and Diverter, FanControl Linkage

18 Screw L

Figure 8−13. Fan Assembly Access Panels




8−5. SERVICING

General

Servicing includes replenishment of fuel, changing or replenishment of oil, andother such maintenance functions (Ref. RMM, Section 12−00−00).

The locations of servicing points are shown in Figure 8−14.

Engine, transmission and hydraulic servicing materials and capacities are shownin Table 8−2. A complete listing of servicing materials may be found in the RMM,Section 12−00−00.

F927−0593

5

12

6

78

4

9

1 ECS − AIR CONDITIONER SYSTEM OIL2 ECS − COMPRESSOR REDUCTION GEARBOX OIL3 FUEL SYSTEM − FUEL4 HYDRAULIC SYSTEM − MANIFOLD/RESERVOIR HY-

DRAULIC FLUID5 LANDING GEAR − DAMPER FLUID

6 TRANSMISSION & DRIVE SYSTEM − MAIN TRANSMIS-SION LUBRICATING OIL

7 POWERPLANT − ENGINE LUBRICATING OIL8 POWERPLANT − ENGINE WASH SOLUTION9 ROTOR BRAKE − HYDRAULIC FLUID

Figure 8−14. Servicing Points




Capacities − Fuel System:JET A: 1097 lb; 498 kg; 161.3 U.S. gal; 611L total capacity

1078 lbs; 158.5 U.S. gal; 600L useable

JET B: 1048 lb; 476 kg; 161.3 U.S. gal; 611L total capacity1030 lbs; 158.5 U.S. gal; 600L useable

Table 8−1. Acceptable Fuels

NOTE: For additional information on fuels, refer to Pratt and Whitney 207E MaintenanceManual

FUEL TYPE

SPECIFICATION

USA CANADA UK FRENCH NATO PRCKerosene:Jet A, A−1, A−2**JP8*

ASTM D1655MIL−T−83133

CGSB3.23−M86

AVTURDERD 2453*DERD 2494* AIR 3405D

F−34F−35

RP−3

Wide Cut:Jet BJP4*

ASTM D1655MIL−T−5624

CGSB3.22−M86

AVTAGDERD 2454*DERD 2486* AIR 3407B F−40

High Flash:JP5* MIL−T−5624

CGSB3.GP−24Ma

AVCATDERD 2452*DERD 2498* AIR 3404C

F−43F−44

* Contains fuel system icing inhibitor (FSII). For JP−8, MIL−T−83133C allows two grades. The grade meeting NATO code F−34 has FSIIwhile the grade meeting code F−35 has no FSII without prior agreement.

** For Jet A−2 conforming to CAN/CGSB 3.23−M86 is acceptable for use, provided the restrictions regarding flash and freezing points arestrictly observed.

Table 8−2. Servicing Materials

Specification Material Manufacturer

1. Engine − Total Capacity 1.34 U.S. Gal (1.12 Imp Gal; 5.12 L)NOTE: The mixing of different oil brands is not approved.

MIL−PRF−23699 Aero Shell Turbine Oil 500 Shell Oil Co.50 W. 50th StNew York, NY 10020

Shell Canada Products Ltd.1500 Don Mills RoadDon Mills, OntarioCanada M3B 3K4





MIL−PRF−23699 Castrol 5000 Castrol Canada, Inc.3660 Lakeshore Blvd. WestToronto, OntarioCanada M8W 1P2

Castrol Specialty Products Div.16715 Von Karman Ave.Suite 230Irvine, CA 92714−4918

Castrol (U.K.) Ltd.Burmah HousePipers WaySwindon, BerkshireSN3 1RE England

Exxon Turbo Oil 2380Exxon Turbo Oil 2525

Exxon International Co.200 Park Avenue Florham Park, NJ 07932−1002

Esso Petroleum Canada55 St. Clair Avenue WestToronto, OntarioCanada M5W 2J8

Exxon Co.P.O. Box 2180Houston, TX 77001

Mobil Jet Oil II Mobil Oil Corp.International Aviation Division150n East 42nd StreetNew York, NY 10017, USA

Mobil Oil Corp.Aviation and Government Sales3225 Gallows RoadFairfax, VA 22037

Esso Petroleum Canada55 St. Clair Avenue WestToronto, OntarioCanada M5W 2J8




Specification ManufacturerMaterial

MIL−PRF−23699 Royco Turbine Oil 500 Royal Lubricants Co. Inc.P.O. Box 518Hanover, NJ 07936

Turbonycoil 525−2A Nyco S.A.66Ave. Des Champs ElyseeParis, France 75008

2. Main Transmission − Total Capacity 10.0 Quarts, 9.5 L

NOTE: Observe servicing instruction placard located on transmission oil filler.

Transmissions P/N 900D1400004−101 and 900D1400005−101:

MIL−PRF−23699 See item 1. Engine

Mobil Jet Oil 254

Transmission P/N 900D1400006−101:

Mobil SHC 626

3. Hydraulic System

MIL−PRF−83282

4. Rotor Brake

MIL−PRF−83282

Fuel system:

Fuel System Servicing Precautions

Only qualified authorized personnel may fuel the helicopter.

Static producing clothing shall not be worn.

Open flames and smoking are not permitted in refueling area.

Refueling vehicle should be parked a minimum of 20 feet from helicopterduring fueling operation.

At least one fully−charged 50 pound CO2 fire extinguisher shall be in theimmediate area.

Before starting fueling operation ground helicopter if possible.

Service fuel cell slowly.




Fuel system filling

NOTE: With the fuel system ‘‘topped off’’, the fuel quantity indication will not displayactual fuel weight. The pilot must visually determine fuel quantity by removingthe fuel cap and noting fuel level on the inside of filler neck (Ref. Figure 8−15).

Fuel helicopter with correct fuel as soon as possible after landing to preventmoisture condensation.

Keep fuel nozzle free of all foreign matter.

Always ground fueling nozzle or fuel truck to GROUND HERE receptacleor to another bare metal location before removing service cap.

Remove the filler cap and secure the lanyard in the slot provided in the fillercap adapter.

NOTE: The lanyard must be secured properly in order to assure that the gravity filler portcheck valve fully opens.

Do not attempt to refuel helicopter if the lanyard has broken.

Service fuel cell slowly.

Secure filler cap after fueling.

Remove fuel nozzle and ground(s) from helicopter.

FUEL SYSTEM FILLER PORT

FUEL CAP LANYARD

F92−117

156 − GAL

152 − GAL

146 − GAL

FUEL QUANTITY MARKS

Figure 8−15. Fuel System Gravity Filler Port

CAUTION




Hydraulic System:

NOTE: The hydraulic system may be serviced by using either the optional hydraulicsystem hand pump or a hydraulic mule. For servicing the system using the mule,refer to the RMM, Section 29−00−00.If the hand pump is not installed, the hydraulic fluid level must monitored closelyand serviced before leaving an area where proper facilities are located.

The hand pump (optional) provides capability to pump fluid into the manifold/res-ervoir without the need of a ground support unit. The hand pump is mountednext to the GSE panels on the transmission deck.

A sight glass indicates when the fluid is at the 0.3 qt level. On the undersideof the cover, a can opener provides a clean means of opening new cans of hydraulicfluid.

A manually operated selector valve is mounted internally in the housing. Theselector valve lever provides for selection of system 1 or system 2 servicing.

The drive handle folds and clips against the reservoir housing for storage. Whenin use, the handle extends through the open access panel, providing a convenientmeans of operation.

NOTE: The following servicing procedure applies to aircraft equipped with the optionalhydraulic system hand pump.

Do not mix different specification hydraulic fluids. Ensure that onlyMIL−H−83282 fluid is used to service the hydraulic systems forall helicopter operations in temperatures above −40°F.The intentional mixing of approved hydraulic oils is not permitted.

Servicing − Hydraulic hand pump:

Open transmission access panel (Ref. Figure 8−6 and Figure 8−7).

Verify that hydraulic fluid is low by checking oil level on hand pump reservoirfluid level sight gauge (Ref. Figure 8−16).

Unscrew the reservoir cover to remove.

Add appropriate amount of hydraulic oil.

Replace cover.

Servicing − Hydraulic system:

Verify that pump reservoir has fluid; replenish if necessary.

Select system to be serviced by using the selector valve lever on the handpump (Ref. Figure 8−16).

Disengage handle from stowed position.

Rotate handle in direction of arrow (CW).

Servicing is complete when the hydraulic manifold fluid level sight gaugeis at the correct level.

Stow handle.

CAUTION




F927−025

MINIMUM OIL LEVEL

HANDLE (STOWED)

HANDLE IN OPERATING POSITION

SELECTOR VALVELEVER

HAND PUMP (OPTIONAL)

HYDRAULIC MAINFOLD/RESERVOIR

FLUID LEVELSIGHT GAUGE

COVER

FLUID TEMP

−40°C 95°CFULL

REFILL

EMPTY

FULL

VIEW LOOKING DOWN

REFILL

Figure 8−16. Hydraulic System




Main Transmission Filling

Open access panel (Ref. Figure 8−6 and Figure 8−7).

Open oil filler cap (Ref. Figure 8−17).

Pour in oil.

Verify quantity of oil in sight window.

NOTE: Correct oil livel is when the observed level is halfway between the “ADD” and“FULL” marks (Ref. Figure 8−17).

Close oil filler cap.

Close access panel (Ref. Figure 8−6 and Figure 8−7).

Main Transmission Draining

Open access panel (Ref. Figure 8−6 and Figure 8−7).

Remove chip detector (Ref. RMM, Section 63−20−00).

Using transmission drain line, place free end of drain line in a suitable container.

Insert drain line probe in chip detector housing.

Allow transmission to drain.

Remove drain line and install chip detector (Ref. RMM, Section 63−20−00).

Close access panel (Ref. Figure 8−6 and Figure 8−7).




CHIP DETECTOR

ADD

FULL

ADD

FULL

VIEW ROTATED

OIL LEVELINDICATOR

F92−119B

OIL FILLER

CHIPDETECTORHOUSING

FILTER BYPASSINDICATOR

TRANSMISSIONOIL FILL

SERVICE WITHMOBIL SHC 626 OIL

NOTE: 900D1400006−101 TRANSMISSION ONLY

TRANSMISSIONOIL FILL

SERVICE WITH OILPER MIL−L−23699

NOTE: 900D1400004−101 AND 900D1400005−101TRANSMISSIONS ONLY

TRANSMISSION IS SERVICEDPROPERLY WHEN OIL LEVEL IS

HALFWAY BETWEEN �FULL" AND �ADD".

Figure 8−17. Main Transmission Servicing




Rotor brake:

The rotor brake reservoir is located on the top forward deck.

Open right−hand forward access door.

Remove filler cap.

Using hydraulic fluid, fill reservoir to top of sight glass.

Install filler cap.

Close right−hand forward access door.

FILLER CAP

SIGHT GLASS

F92−120

Figure 8−18. Rotor Brake




Powerplant

Engine Oil System Filling / Replenishing

Do not mix different brands or types of oil since their differentchemical structures may make them incompatible. If differentbrands or types of oil become mixed, drain system (includingengine integral oil tank, engine oil filter housing, engine oil heatexchanger and oil in and out hoses) and refill with new oil.

NOTE: To reduce the possibility of over filling the oil tank, check the oil level 10 minutesafter engine shutdown.

Open oil dipstick hand hold/door (Ref. Figure 8−6 and Figure 8−7).

Remove engine oil filler cap.

Refill engine oil tank with specified oil in related manufacturers’ publications(Ref. RMM, Section 01−00−00)

NOTE: Correct oil level is when the observed level is between the MAX and MIN markson the oil dipstick. Filling the oil tank to MAX may result in oil being ventedoverboard, causing a buildup of carbon deposits on the tailboom andempennage. Should this occur, monitor engine oil level without adding oil (unlessthe oil level falls below MIN) to determine if the level stabilizes at some pointbetween MAX and MIN. Once this level is determined, fill oil to and maintain thislevel.

Replace oil filler cap.

Install and lock the oil filler cap on the oil transfer tube as follows (Ref.Figure 8−19).

Make sure to install the oil filler cap correctly. Incorrect installationcan lead to disengagement of the cap locking lugs; the cap canthen lift from its locking position and have an incorrect sealing.This can result in an oil loss that may require shuting down theengine.

Place the dipstick in the gearbox and make sure that the dipstick off−setof the cap is in line with the off−set hole of the oil filler tube of the gearbox.Turn the handle and lock the cap. Make sure that the cap handle is in thelock position.

If extra force is required to lock the cap, it means that the capis not installed correctly. Remove the cap and reinstall it.

NOTE: The writing on the cap handle should be facing toward the front of the engine.

Close oil dipstick hand hold/door.

CAUTION

CAUTION

CAUTION




VIEW LOOKING AFT

CHIP DETECTOR

OIL FILTER IMPENDINGBYPASS INDICATOR

F92−121A

OIL FILLER DIPSTICK(VIEW ROTATED)

NOTE: SOME ENGINE DETAIL OMITTED FOR CLARITY

WRITING ON TAB FACING FORWARD

Figure 8−19. Engine Oil System − Servicing




8−6. AIRCRAFT CLEANING

General cleaning of oil and dirt deposits from the helicopter and its componentscan be accomplished by using dry−cleaning solvent, standard commercial gradekerosene, or a solution of detergent soap and water. Exceptions that must be observedare specified in the following cleaning paragraphs.

Storage, use, and disposal of all solvents must be per Governmentand local health and safety regulations.

Fuselage Interior Trim and Upholstery

Fuselage Interior Trim and Upholstery Cleaning

Carpet cleaning agents may damage underlying metal or compositesurfaces. Carpet or seats must be removed from helicopter priorto cleaning and allowed to air dry prior to reinstallation.

Clean dirt or dust accumulations from floors and other metal surfaces withvacuum cleaner or small hand brush.

Any flammable solvent that may affect material flammability mustbe removed completely after cleaning.

Sponge soiled upholstery and trim panels with a mild soap and lukewarmwater solution. Avoid complete soaking of upholstery and trim panels. Wipesolution residue from upholstery with cloth dampened by clean water.

Use solvents sparingly. Some solvents may soften or dull material.Test an inconspicuous area prior to use.

Remove imbedded grease or dirt from upholstery and carpeting by spongingor wiping with an upholstery cleaning solvent.

Helicopter Exterior

Main Rotor Blade Cleaning

Use care to prevent scratching of fiberglass skin when cleaningmain rotor blades. Never use volatile solvents or abrasive materials.Never apply bending loads to blades or blade tabs during cleaning.

NOTE: Avoid directing high pressure concentrations of soap and/or clean water towardengine air intake areas, instrument static source ports and main rotor swashplatebearings.

Clean rotor blades when necessary using solution of clean water and mildsoap.

CAUTION

CAUTION

CAUTION

CAUTION

CAUTION




Fuselage Exterior Cleaning

NOTE: Avoid directing high pressure concentrations of soap and/or clean water towardengine air intake areas, instrument static source ports and main rotor swashplatebearings.

NOTE: Check and drain, if moisture present, the static system drain valves after theaircraft has been washed or exposed to rain or snow and any time the airspeedor altimeter indicators are showing sporadic readings. (Ref. Figure 8−6 andFigure 8−7).

Clean helicopter exterior, including fiberglass/kevlar components, when nec-essary, use solution of clean water and mild soap.

Transparent Plastic

Transparent Plastic Cleaning

Clean outside surfaces of plastic panels by rinsing with clean water and rub-bing lightly with palm of hand.

Use mild soap and water solution or aircraft type plastic cleaner to removeoil spots and similar residue.

Never attempt to dry plastic panels with cloth. To do so causesany abrasive particles lying on plastic to scratch or dull surface.Wiping with dry cloth also builds up an electrostatic charge thatattracts dust particles from air.

After dirt is removed from surface of plastic, rinse with clean water and letair−dry.

Clean inside surfaces of plastic panels by using aircraft type plastic cleanerand tissue quality paper wipers.

8−7. COCKPIT DOOR REMOVAL

Disengage gas strut from cockpit floor attachment (Ref. Figure 8−20).

NOTE: Fit between socket end of strut and ball end of ball stud is by interference.Removal of strut from its attachment requires a snap action motion to pull awaythe socket end from the ball stud.

Remove lower quick release pin by pulling on the ring.Remove door restraint by pulling away from lower fork assembly.Remove upper quick release pin while holding the door.Remove door.Install quick release pins into upper and lower fork assemblies.

Installation is opposite of removal.

CAUTION




F92−123A

BALL SOCKET BALL STUD

VIEW LOOKING OUTBOARDLEFT SIDE SHOWN, RIGHT SIDE OPPOSITE

DOOR RESTRAINT

DOORFRAME

WINDOWFRAMELOWER DOOR ATTACHMENT WITH

REMOVABLE RESTRAINT. UPPERDOOR ATTACHMENT SIMILAR.

GAS STRUT

QUICKRELEASE PIN

RING

LOWERFORKASSEMBLY

Figure 8−20. Cockpit Door Attachment




8−8. CABIN SEATS: REMOVAL/INSTALLATION

Cabin Seat Removal (Ref. Figure 8−21):

First disengage upper quick disconnect fittings then lower quick disconnect fit-tings from their anchor plates to release cabin seat assembly. Remove seat assem-bly.

Cabin Seat Installation (Ref. Figure 8−21):

Align cabin seat assembly with floor anchor plates. First engage lower quickdisconnect fittings then upper quick disconnect fittings of cabin seat assemblyinto mating roof and floor anchor plates to secure.Ensure fittings are fully and properly engaged.

ANCHOR PLATE

LOWER SEAT ATTACHMENT

UPPER SEAT ATTACHMENT

F92−124

QUICK DISCONNECT

NOTE: PULLING/RELEASING THE KNURLED COLLAR ONTHE QUICK DISCONNECT RELEASES/ENGAGESTHE LOCKING MECHANISM

KNURLED COLLAR

ANCHOR PLATE

Figure 8−21. Cabin Passenger Seat Attachment




8−9. COPILOT FLIGHT CONTROLS

F92−125

PEDAL DISENGAGE PIN

PEDAL SHAFTASSEMBLY

PEDAL ADJUSTMENT PINS

PEDAL CRANKASSEMBLY

HOOK TAPE

Figure 8−22. Copilot PedalsCopilot Pedals: Disengaging (Ref. Figure 8−22)

Copilots pedal shaft assemblies can be temporarily stowed in the full forwardposition.

Pull up pedal adjustment pins.

Pull out pedal disengage pin.

Swing shaft assemblies forward to their hook tape secured positions.

Copilot Pedals: Engaging

Reengaging copilot’s pedals is opposite of disengaging.

NOTE: Ensure that pedal adjustment pins are fully seated in pedal crank assemblies.




8−10.ENGINE CHARTS

The following charts define maintenance action requirements for engine over temper-ature, overspeed, and overtorque.

980

960

940

920

900

880

870

860

840

820

800

0 20 30 60 90 120 150 180 210 240 270 300 330 SECMIN1 2 3 4 5

AREA C

NO ACTION REQUIRED

AREA A − RECORD IN ENGINE LOG BOOK(2.5 MINUTE RATING)

AREA B − RECORD IN ENGINE LOG BOOK(CONTINUOUS OEI)

AREA C − RETURN ENGINE FOR OVERHAUL

DO AN HSI IF THE FOLLOWING CONDITION OCCURS:15 MINUTES OF CUMULATIVE RUNNING TIME IN AREA A.

TIME (MINUTES AND SECONDS)

EX

HA

US

T G

AS

TE

MP

ER

AT

UR

E

F927−054

930

890

970ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ

AREA A

850

910

950

9901000

ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ

AREA BTAKE−OFF(FIVE MINUTES)

AREA C

Figure 8−23. EGT Vs Time − All Conditions Except Starting




820

20 30

AREA C

NO ACTION REQUIRED

AREA A − VISUAL INSPECTION THROUGH EXHAUST DUCT ANDRECORD IN ENGINE LOG BOOK

AREA B − PERFORM HSIAREA C − RETURN ENGINE TO OVERHAULAREA D − DETERMINE CAUSE FOR HUNG START AND CARRY

OUT DRY MOTORING RUN PRIOR TO ATTEMPTINGA RE−START

TIME (SECONDS)

ME

AS

UR

ED

GA

S T

EM

PE

RA

TU

RE

(E

GT

)

2

650

760

875

45

AREA D

ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ

ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ

AREA A

AREA B

F92−127

Figure 8−24. EGT Vs Time − Starting

0 20 30 40 50 SEC

NO ACTION REQUIRED

AREA A − RETURN ENGINE TO OVERHAUL

TIME (SECONDS)

PO

WE

R T

UR

BIN

E S

PE

ED

(%

RP

M)

100.0

104.5

112.4

10

F92−128

Figure 8−25. Power Turbine (NP) Speed Vs Time




ÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉÉ

0 20 40 60 80 120 SECMIN1 2

AREA A − RECORD IN ENGINE LOG BOOK (2.5 MINUTE RATING)

AREA B − RECORD IN ENGINE LOG BOOK (CONTINUOUS RATING)

AREA C − RETURN ENGINE FOR OVERHAUL


GA

S G

EN

ER

AT

OR

SP

EE

D −

%

NG

100 1402.5

100.0

ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ

AREA C103.0

F927−055

98.2

160 180 200 220 240 260 280 3003 4 5

104.1

AREA B

AREA A

97.2NO ACTION REQUIRED

99.8

TAKE−OFF(5 MINUTES)

Figure 8−26. Compressor Turbine (NG) Speed Vs Time

ÑÑÑÑÑÑ

5MIN

ÑÑÑÑÑÑÑÑÑ

2.5MIN

ÓÓÓÓÓÓ

ÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇÇ ÉÉÉÉ

ÉÉÉÉÉÉÉÉ

F927−056A

AREA C

AREA A

AREA B

020

30 60 90 120 150 180 210 240 270 300 330 (SEC)

1 2 2.5 3 4 5 (MIN)

100%

88.4%

116.7

112.7%

AREA A DETERMINE CAUSE AND RECORD IN ENGINE LOG BOOK

AREA B RETURN ENGINE FOR OVERHAUL

AREA C RECORD IN ENGINE LOG BOOK EXCEPT OEI CONDITION


TORQUEIIDS (XMSN) ENGINE

NOTE: THE TORQUE DISPLAYED ONTHE IIDS PRIMARY DISPLAYIS TRANSMISSION TORQUE.

110%

124%

135%

145%

OEIAEO

100%

MCP

MCP

10SEC

140%

Figure 8−27. Engine Overtorque Limits − All Conditions




8−11.SPECIAL OPERATIONAL CHECKS AND PROCEDURES

The following checks are typically utilized as part of a post maintenance run upassociated with scheduled inspections, troubleshooting or maintenance on specificaircraft systems. Refer to the Rotorcraft Maintenance Manual (CSP900RMM−2)for the recommended use and frequency of the following checks.

These checks and or procedures are to be performed only whileaircraft is on the ground.

ENGINE NP OVERSPEED TEST PROCEDURE

NOTE: This procedure is to be performed with one engine at idle and the other engineoff.

� Engine control switch IDLE� Collective INCREASE TO 20% TORQUE� OVSP test switch MOVE TOWARD OPERATING ENGINE AND HOLD� � Primary IIDS display OBSERVE DECREASE IN TORQUE AND NR� OVSP test switch RELEASE� � Primary IIDS display OBSERVE INCREASE IN TORQUE AND NR� Collective FULL DOWN� Engine control switch OFF� Start other engine and repeat procedure.

ENGINE OUT/LOW ROTOR WARNING CHECK

� Engine control switches FLY

� Collective INCREASE TO A CLP OF >5% FOR MORETHAN 5 SECONDS

� Collective LOWER TO A CLP OF 0%

� Engine control switches SWITCH TO IDLE AND NOTE THE FOLLOWING

� ENG OUT warning light flashes and low rotor RPM tone is activated for one cycle. Asrotor RPM decreases through 88%, the low rotor RPM tone will reactivate until theAOG logic disables the warning.

CAUTION




HYDRAULIC SYSTEM CHECK

� With the aircraft operating at IDLE or FLY (100%):

� � IIDS secondary display CHECK ‘‘1 HYD 2” CAUTION SEGMENTS ARE NOTILLUMINATED

� With the aircraft operating at FLY (100%):

� � HYD TEST switch SET AND HOLD TO ‘‘SYS 1”

� � IIDS secondary display CHECK ‘‘1 HYD” CAUTION SEGMENT ILLUMINATES

� � IIDS alphanumeric display VERIFY 250 PSI MAXIMUM FOR HYD 1 AND 1,000+100/−50 PSI FOR HYD 2

� � HYD TEST switch RELEASE AND CHECK ‘‘1 HYD” CAUTIONSEGMENT OFF

� � HYD TEST switch SET AND HOLD TO ‘‘SYS 2”

� � IIDS secondary display CHECK ‘‘HYD 2” CAUTION SEGMENT ILLUMINATES

� � IIDS alphanumeric display VERIFY 250 PSI MAXIMUM FOR HYD 2 AND 1,000+100/−50 PSI FOR HYD 1

� � HYD TEST switch RELEASE AND CHECK ‘‘HYD 2” CAUTIONSEGMENT OFF

VSCS CHECK

NOTE: This functional check may be performed with the engines off and aircraftconnected to an external power source.

� Left and right VSCS switches OFF� IIDS alphanumeric display VERIFY CAUTION SEGMENT ON AND ‘‘TOTAL STAB

FAIL” INDICATION� VSCS indicator needles CENTERED� Left VSCS switch ON� Right VSCS switch MOMENTARILY TO ‘‘TEST” AND THEN TO ‘‘ON’’� IIDS alphanumeric display VERIFY ‘‘RIGHT STAB FAIL” INDICATION FOR 5 TO 8

SECONDS, THEN OUT� Left VSCS switch MOMENTARILY TO ‘‘TEST” AND THEN TO ‘‘ON’’� IIDS alphanumeric display VERIFY ‘‘LEFT STAB FAIL” INDICATION FOR 5 TO 8

SECONDS, THEN OUT� VSCS indicator needles VERIFY NEEDLES ARE APPROXIMATELY 55% RIGHT

OF CENTER WITH 0% CLP AND AIRCRAFT LEVEL

NOTE: If the selected VSCS system fails the test, the failure annunciation will remainon the IIDS alphanumeric display.




WET ENGINE MOTORING RUN

Before performing this procedure, insure that the power supplyto the ignition exciter is disconnected (IGNTR circuit breakerspulled).

When a fuel metering unit/pump is replaced in the field, motoringor starting the engine is not recommended until priming isaccomplished by performing a engine wet motoring run.

� Twistgrip on selected engine(s) NORMAL

� Fuel valve ON CHECK

� Fuel boost pump ON

� Engine control switch IDLE

NOTE: Maintain starter operation for desired duration while observing starter limits.

� Engine control switch OFF

� Fuel boost pump OFF

NOTE: After a wet motoring run, a dry motoring should be accomplished before any startis attempted.

DRY ENGINE MOTORING RUN

NOTE: This procedure is used to clear internally trapped fuel and vapor from the engine.This procedure maybe used if there is evidence of a fire within the engine or lackof EGT indication after lightoff at the beginning of an engine start.

� Twist grip OFF� Engine control switch for selected engine SET TO IDLE −

OBSERVE STARTER TIME LIMITS� Engine control switch for selected engine OFF

� EEC RESET button PRESS

� Twistgrip PLACE IN NORMAL DETENT

� EEC MAN or flashing indication CHECK OFF

CAUTION

CAUTION




ENGINE WASH PROCEDURES

Engine Water Wash − Desalination

Open main transmission access door (Ref. Figure 8−6 and Figure 8−7).

NOTE: If cleaning agent is to be used, prepare solution and compressor wash systemin accordance with related manufacturers’ publications (Ref. RMM, Section01−00−00)

Use of correct mixture as specified in the PWC Maintenance manualis very important, not only when the temperature is below freezingat the time of washing, but also if the temperature is expected tobe below 2°C (36°F) between time of washing and the next start.

Connect cleaning solution or water source to engine wash panel using ANtype fittings.

NOTE: To prevent precipitation of deposits through the use of hard water, engine mustbe allowed to cool to below 65°C (150°F). Minimum cooling period of 40 minutesmust be allowed since the engine was last operated.

Ensure inlet particle separator and heat / defog shutoff valves areturned off.

Do not motor engine for more than 30 seconds.

NOTE: Ensure cleaning solution or water source pressure of 60−82 PSI.

Perform dry engine motoring run; when NG reaches 5%, inject water solutioninto air inlet case.

Close tank valve as soon as NG falls to 5%.

Allow starter to cool between runs.

If water/methanol mixture has been used, perform additional dry engine mo-toring run.

Close main transmission access door (Ref. RMM, Section 06−00−00).

Repeat procedure on other engine.

Once engine wash is complete, start and operate engines at idle for at leastone minute to completely dry engines.

CAUTION

CAUTION

CAUTION




Engine Wash − Performance Recovery

Open main transmission access door (Ref. Figure 8−6 and Figure 8−7).

NOTE: If cleaning agent is to be used, prepare solution and compressor wash systemin accordance with related manufacturers’ publications (Ref. RMM, Section01−00−00)

Use of correct mixture as specified in the PWC Maintenance manualis very important, not only when the temperature is below freezingat the time of washing, but also if the temperature is expected tobe below 2°C (36°F) between time of washing and the next start.

Connect cleaning solution or water source to engine wash panel using ANtype fittings.

NOTE: To prevent precipitation of deposits through the use of hard water, engine mustbe allowed to cool to below 65°C (150°F). Minimum cooling period of 40 minutesmust be allowed since the engine was last operated.

Ensure inlet particle separator (IPS) and heat / defog shutoff valvesare turned off.Do not motor engine for more than 30 seconds.

NOTE: Ensure cleaning solution or water source pressure of 60−82 PSI.

Perform dry engine motoring run; when NG reaches 5%, inject wash solutioninto air inlet case.

Close tank valve as soon as NG falls to 5%.

Allow cleaning solution to soak for 10 minutes.

Perform dry engine motoring run; when NG reaches 5%, inject one half ofrinse solution into air inlet case.

Observe starter cooling period.

Perform dry engine motoring run; when NG reaches 5%, inject remainderof rinse solution into air inlet case.

If water/methanol mixture has been used, perform second dry engine motoringrun.

Close main transmission access door.

Repeat procedure on other engine.

Once engine wash is complete, start and operate engines at idle for at leastone minute to completely dry engines.

CAUTION

CAUTION




RIGHT ENGINE WASHTUBE ASSEMBLY

ENGINE WASHNOZZLE

F92−122

VIEW ROTATED

ENGINEWASH

LEFT

RIGHT

LEFT ENGINE WASHTUBE ASSEMBLY

Figure 8−28. Engine Wash Panel




MANUAL ENGINE SHUTDOWN CHECK

NOTE: This procedure should be performed with engine control switch in IDLE and allunnecessary bleed air and electrical equipment, including generator, OFF.

� Twist grip IDLE DETENT

� NP slows to idle CHECK

� EEC MAN indication on primary IIDS dis-play

CHECK

� Twistgrip SNAP TO CUTOFF

� Engine control switch OFF

� IIDS CHECK NORMAL SHUTDOWNINDICATIONS

� NG zero percent CHECK� EEC RESET button PRESS

� Twistgrip PLACE IN NORMAL DETENT

DO NOT return twist grip to the NORMAL detent until NG is at zeroand the EEC RESET button is pressed. Failure to follow thisprocedure may cause a re−light with a subsequent EGTexceedance.

� EEC MAN indication CHECK OFF

CAUTION




MANUAL ENGINE START CHECK

NOTE: Complete the Engine Prestart cockpit check (Ref. Section IV) before attemptinga manual start.

� Twistgrip ROTATE TO FULL OPEN (PASTTHE ‘‘NORMAL’’ DETENT)

� EEC MAN indication on primary IIDS dis-play

CHECK

� Twist grip ROTATE TO OFF

� Generator OFF

� L BOOST or R BOOST ON, CHECK IIDS INDICATIONS

� Engine control switch IDLE

� Twist grip ROTATE TOWARD IDLE AS NGINCREASES THROUGH 8PERCENT

NOTE: As NG increases through 8% rotate twistgrip toward normal until lightoff occurs.Observe EGT indication for immediate temperature rise. Monitor EGT and NGduring start. Observe start limits. Increase twistgrip toward normal only asnecessary to keep NG accelerating toward idle. Manually bring NP/NR to 65%.

If lightoff is not attained with an increase of EGT and NG within10 seconds, rotate the twistgrip to OFF and place the engine controlswitch to off. Following a 30 second fuel drain period, perform a30 second dry motoring run before attempting another start. Repeatthe complete starting sequence observing limitations.

� EEC RESET button PRESS WHEN NP/NR IS 65PERCENT

� EEC MAN indication CHECK FLASHING� Twistgrip NORMAL DETENT

� EEC MAN indication CHECK OFF

� Engine oil pressure CHECK

� Generator ON

� IIDS CHECK

CAUTION




AUTOROTATION RPM CHECK

Refer to CSP−900RMM−2, Section 18−00−00.

NOTE: This procedure should be performed with engine control switches in FLY andcollective full down. However, aircraft operating at or near gross weight limits andat high density altitudes may not be able to perform this procedure with collectivefull down without exceeding rotor limits. Refer to CSP−900RMM−2, Section18−00−00 for alternative collective position while operating at high grossweights.

� Target altitude SELECT

NOTE: Select an altitude above target altitude so as to arrive at the target altitude insteady state autorotation at 70 KIAS. Failure to maintain constant airspeedduring autorotation will cause rotor RPM fluctuations, resulting in inaccurateRPM readings.

� IIDS SELECT “CLP” ONALPHANEUMERIC DISPLAY

� Airspeed 70 KIAS� Collective lever position ZERO % OR 10% AS REQUIRED

Observe rotor limits.

� At target altitude RECORD ROTOR RPM

NOTE: If gross weight/density altitude combination allows procedure with collective fulldown, the torque reading should be zero percent at target altitude for accurateautorotation RPM.

BATTERY REMOVAL

The battery should be removed from the helicopter and placed in a heated premisesif ambient air temperatures of minus 18°C or less or when long time exposure tocold temperatures is expected.

CAUTION




RESETTING IIDS TIME/DATE

SET TIME/DATE


NOTE 1

F92−102


TIME <HH:MM>

DATE MM−DD−YYNOTE 1.: ‘‘ENT’’ KEY SELECTS FIELD TO BE SET (MINUTE,

HOURS, DAY, MONTH, YEAR) AND SELECTED FIELDBLINKS, � AND � KEYS INCREMENT/DEINCRE-MENT DIGIT VALUE, ‘‘REC’’ KEY CHANGES TIMEAND DATE TO SELECTED VALUES

Figure 8−29. Set Time/DateTo change date or time:

This procedure is to be perfromed with both engines OFF.

Press MENU to enter menu system.

Use ↑ or ↓ keys to select SET TIME/DATE and press ENT. The following isdisplayed on the IIDS two line alphanumeric display:TIME HH:MMDATE MM−DD−YY

Press ENT to edit display. The hour digits in the TIME HH:MM display willblink indicating these are the digits selected for editing.

Use ↑ or ↓ keys to change value of flashing digit/value.

NOTE: Holding the arrow key for more than one second will cause the value of thedigit(s) being edited to increment at the rate of one per second.

Press ENT to select next digit(s) (the minutes digits will blink) and set valueusing ↑ or ↓ keys.

NOTE: Each press of the ENT key will select the next value to edit in the sequence theyare displayed.

Repeat above steps until the correct time and date is displayed.

Use the REC key to save the changed time/date. Pressing the CLR key insteadof REC will abandon all changes.

NOTE: The REC key may be pressed at any time during the editing process to save thechanges made. Any fields not changes will remain at their present values.

Additional Operationsand Performance Data


Original 9−i/( 9−ii blank)

Reissue 1

S E C T I O N I XADDITIONAL OPERATIONSAND PERFORMANCE DATA

TABLE OF CONTENTS

PARAGRAPH PAGE9−1. Abbreviated Checklists 9−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

9−2. Fuel Flow vs Airspeed 9−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 9−1. Fuel Flow, AEO, Sea Level, ISA (15°C) 9−5. . . . . . . . . . . . . . . . . . . . . . .

Figure 9−2. Fuel Flow, AEO, 4000 Feet HP , ISA (7°C) 9−6. . . . . . . . . . . . . . . . . . . .

Figure 9−3. Fuel Flow, AEO, 8000 Feet HP, ISA (−1°C) 9−7. . . . . . . . . . . . . . . . . . . .

Figure 9−4. Fuel Flow, AEO, Sea Level, ISA + 20°C (35°C) 9−8. . . . . . . . . . . . . . . . .

Figure 9−5. Fuel Flow, AEO, 4000 Feet HP, ISA + 20°C (27°C) 9−9. . . . . . . . . . . . .


Figure 9−7. Fuel Flow, AEO, Sea Level, ISA + 30°C (45°C) 9−11. . . . . . . . . . . . . . . . .



Figure 9−10. Fuel Flow, OEI, Sea Level, ISA (15°C) 9−14. . . . . . . . . . . . . . . . . . . . . . .

Figure 9−11. Fuel Flow, OEI, 4000 Feet HP , ISA (7°C) 9−15. . . . . . . . . . . . . . . . . . . .

Figure 9−12. Fuel Flow, OEI, 8000 Feet HP , ISA (−1°C) 9−16. . . . . . . . . . . . . . . . . . .

Figure 9−13. Fuel Flow, OEI, Sea Level, ISA + 20°C (35°C) 9−17. . . . . . . . . . . . . . . .

Figure 9−14. Fuel Flow, OEI, 4000 Feet HP , ISA + 20°C (27°C) 9−18. . . . . . . . . . . .


Figure 9−16. Fuel Flow, OEI, Sea Level, ISA + 30°C (45°C) 9−20. . . . . . . . . . . . . . . .



Figure 9−19. Fuel Flow, OEI, −1000 Feet HP , ISA (17°C) 9−23. . . . . . . . . . . . . . . . . .

Figure 9−20. Fuel Flow, OEI, −1000 Feet HP , ISA + 20°C (37°C) 9−24. . . . . . . . . . .

Figure 9−21. Fuel Flow, OEI, −1000 Feet HP , ISA + 30°C (47°C) 9−25. . . . . . . . . . .

9−3. International Civil Aviation Organization (ICAO) NoiseLevels 9−26. . . . . . . . . . . . . .


MD900 (902 Configuration with PW 207E) Additional Operationsand Performance Data


SECTION IXADDITIONAL OPERATIONS AND

PERFORMANCE DATA

9−1. ABBREVIATED CHECKLISTS

NOTE: These checklists do not have any CAUTION, WARNINGS, or NOTES. Be sureyou have a thorough understanding of the checks as described in Section IVbefore attempting to operate the helicopter.

ENGINE PRE−START COCKPIT CHECK

ELECTRICAL POWER − OFF

� All cabin doors CHECK

� Seat belt and shoulder harness FASTENED

� Rotor brake STOWED

� Flight instruments CHECK STATIC POSITION/SET

� Collective friction ON

� Collective stick position FULL DOWN

� Twistgrip alignment marks aligned with in-dex mark

CHECK

� LDG/HVR lights OFF

� Key switch ON

� Circuit breakers IN

� Utility panel switches OFF EXCEPT VSCS ON

� NACA inlet switch AS REQUIRED

� Lighting control panel switches AS REQUIRED

� Avionics AS DESIRED

� L GEN and R GEN ON (OFF FOR GPU START)

� POWER OFF

� L BOOST AND R BOOST OFF

� LEFT/RIGHT FUEL SHUTOFF ON; COVER CLOSED

� L ENGINE and R ENGINE OFF


MD900 (902 Configuration with PW 207E)Additional Operationsand Performance Data


ELECTRICAL POWER − ON

� POWER BAT/EXT

� Monitor BIT FIRE WARNING ANNUNCIATORS ONFOR 2 SECONDS; CHECK IIDS FORADVISORIES AT COMPLETION OF BIT

� Fuel quantity display CHECK

� DISP (display by exception) AS DESIRED

ENGINE STARTING − AUTOMATIC

� L BOOST or R BOOST ON; CHECK IIDS INDICATIONS

● EEC MAN indicators OFF

� L ENGINE or R ENGINE SET TO IDLE/FLY AS REQUIRED

� IIDS CHECK FOR NORMALINDICATIONS

� Repeat starting procedure for second engine

� GPU start only:

● L GEN/R GEN ON

● GPU DISCONNECT

ENGINE RUNUP

� Avionics ON, AS DESIRED

� L ENGINE and R ENGINE FLY




BEFORE TAKEOFF

� Cyclic control CHECK RESPONSE

� Collective friction AS DESIRED

� Primary and secondary IIDS displays CHECK ADVISORIES

� Utility panel switches AS REQUIRED

ENGINE/AIRCRAFT SHUTDOWN − NORMAL

� Collective stick FULL DOWN/FRICTION ON

� Cyclic stick TRIM TO NEUTRAL

� Pedals NEUTRAL

� L ENGINE and R ENGINE IDLE

� All unnecessary electrical equipment OFF

� Heat OFF

� AC (if installed) OFF

� Pitot heat (if installed) OFF

� IPS (if installed) OFF

� Lighting control panel AS DESIRED

� Avionics master switch OFF

� L GEN/R GEN switches OFF

� L BOOST/R BOOST OFF

� L ENGINE and R ENGINE OFF

� ENG OUT indications CHECK IIDS FOR NORMALINDICATIONS

� Rotor brake (if installed) APPLY BELOW70% NR

� IIDS CHECK FOR INDICATIONSOR MESSAGES

� POWER OFF




9−2. FUEL FLOW vs AIRSPEED

Description: The fuel flow charts presented in this section are based on level flightperformance data. Fuel consumption values are based on minimum specificationengines and thus may vary between engines. This data is based on a baseline aircraftwith 15% electrical load, engine bleeds and air conditioner off.

Use of Chart: Use the charts as illustrated by the example below.

NOTE: The following example uses Figure 9−1.

Example:

Wanted: Rate of fuel flow

Known: Airspeed = 115 KIAS

Known: Estimated gross weight = 5500 pounds

Method: Enter the chart at the known airspeed of 115 knots (interpolation re-quired). Move vertically to the 5500 pound point (interpolation required)then move to the left to the fuel flow scale and read a fuel flow of approxi-mately 460 LB/HR.




F927−026−1B

INDICATED AIR SPEED − KNOTS

FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

450

500

550

600

40 50 60 70 80 90 100 110 120 130 140 150

Long range cruise

6,000 lb.

5,000 lb.

4,000 lb.

6250 lb.

6500 lb.

MCP limit

Figure 9−1. Fuel Flow, AEO, Sea Level, ISA (15°C)




F927−026−2A

6,000 lb.

5,000 lb.

4,000 lb.

MCP limit

6250 lb.

Long range cruise

6500 lb.

200

250

300

350

400

450

500

550

600


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150

Figure 9−2. Fuel Flow, AEO, 4000 Feet HP , ISA (7°C)




F927−026−3A

6,000 lb.

5,000 lb.

4,000 lb.

MCP limit

6250 lb.

Long range cruise

6500 lb.

200

250

300

350

400

450

500

550


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150

Figure 9−3. Fuel Flow, AEO, 8000 Feet HP, ISA (−1°C)




F927−027−1

200

250

300

350

400

450

500

550


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150

MCP limit

Long range cruise

6,000 lb.

5,000 lb.

4,000 lb.

6250 lb.

6500 lb.

600

Figure 9−4. Fuel Flow, AEO, Sea Level, ISA + 20°C (35°C)




F927−027−2A

MCP limit

Long range cruise

6,000 lb.

5,000 lb.

4,000 lb.

6250 lb.

6500 lb.

200

250

300

350

400

450

500

550


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150

Figure 9−5. Fuel Flow, AEO, 4000 Feet HP, ISA + 20°C (27°C)




F927−027−3A

200

250

300

350

400

450

500


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150

MCP limit

Long range cruise

6500 lb.

6,000 lb.

5,000 lb.

6250 lb.

4,000 lb.





F927−028−1A

MCP limit

6500 lb.

6,000 lb.

5,000 lb.

6250 lb.

4,000 lb.

Long range cruise

200

250

300

350

400

450

500


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150

550

Figure 9−7. Fuel Flow, AEO, Sea Level, ISA + 30°C (45°C)




F927−028−2A

MCP limit

6500 lb.

6,000 lb.

5,000 lb.

6250 lb.

4,000 lb.

Long range cruise

200

250

300

350

400

450

500


FU

EL

FL

OW

− L

B/H

R

40 50 60 70 80 90 100 110 120 130 140 150





F927−028−3

MCP limit

6500 lb.

6,000 lb.

5,000 lb.

6250 lb.

4,000 lb.

Long range cruise


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

450

500

40 50 60 70 80 90 100 110 120 130 140 150





F927−029−1A


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100 110

6500 lb.

6,000 lb.

5,000 lb.

6250 lb.

4,000 lb.

MCP limit

Figure 9−10. Fuel Flow, OEI, Sea Level, ISA (15°C)




F927−029−2A


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100 110

MCP limit

6500 lb.

6,000 lb.

5,000 lb.

6250 lb.

4,000 lb.

Figure 9−11. Fuel Flow, OEI, 4000 Feet HP , ISA (7°C)




F927−029−3A


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100 110

6,000 lb.

5,000 lb.

4,000 lb.

MCP limit

Figure 9−12. Fuel Flow, OEI, 8000 Feet HP , ISA (−1°C)




F927−030−1A


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100 110

MCP limit

6250 lb.

4,000 lb.

6,000 lb.

5,000 lb.

Figure 9−13. Fuel Flow, OEI, Sea Level, ISA + 20°C (35°C)




F927−030−2A

MCP limit

5,000 lb.

4,000 lb.


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100 110

Figure 9−14. Fuel Flow, OEI, 4000 Feet HP , ISA + 20°C (27°C)




F927−030−3A

MCP limit

5,000 lb.

4,000 lb.


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100





F927−031−1A

MCP limit

5,000 lb.

4,000 lb.


FU

EL

FL

OW

− L

B/H

R

200

250

300

350

400

40 50 60 70 80 90 100

Figure 9−16. Fuel Flow, OEI, Sea Level, ISA + 30°C (45°C)




F927−031−2

200

250

300

350

400

40 50 60 70 80 90 100Indicated Airspeed − Knots

Fu

el F

low

− lb

./hr.

MCP limit

5,000 lb.

4,000 lb.





F927−031−3

200

250

300

350

400


Fu

el F

low

− lb

./hr.

MCP limit

4,000 lb.

4500 lb.





F927−032−1

200

250

300

350

400


Fu

el F

low

− lb

./hr.

MCP limit

6,000 lb.

5,000 lb.

4,000 lb.

6250 lb.

Figure 9−19. Fuel Flow, OEI, −1000 Feet HP , ISA (17°C)




F927−032−2

200

250

300

350

400


Fu

el F

low

− lb

./hr.

MCP limit

6,000 lb.

4,000 lb.

6250 lb.

5,000 lb.

Figure 9−20. Fuel Flow, OEI, −1000 Feet HP , ISA + 20°C (37°C)




F927−032−3

200

250

300

350

400

40 50 60 70 80 90 100

Indicated Airspeed − Knots

Fu

el F

low

− lb

./hr.

MCP limit

6,000 lb.

5,000 lb. 4,000 lb.

Figure 9−21. Fuel Flow, OEI, −1000 Feet HP , ISA + 30°C (47°C)




9−3. INTERNATIONAL CIVIL AVIATION ORGANIZATION (ICAO) NOISELEVELS

The MD900 meets the ICAO Annex 16, Volume 1, Chapter 8 noise requirementsfor level flight, takeoff/climb, and approach descent profiles at the certified maximumgross weight of 6250 LB.

MD900 ENGINE: PW 207E GROSS WEIGHT: 6250 LB

Configuration Level FlyoverEPNL

(EPNdB)

TakeoffEPNL

(EPNdB)

ApproachEPNL

(EPNdB)

Clean aircraft, doors on, noexternal kits.

83.50 85.40 89.62


Optional Equipment


S E C T I O N XOPTIONAL EQUIPMENT

TABLE OF CONTENTS

PARAGRAPH PAGE10−1. General Information 10−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−2. Listing − Optional Equipment 10−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 10−1. Optional Equipment MD900 Helicopter 10−2. . . . . . . . . . . . . . . . . . . . . . .

10−3. Compatibility − Combined Optional Equipment 10−2. . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 10−2. Optional Equipment Kit Compatibility − MD900 Helicopter 10−2. . . .

10−4. Optional Equipment Performance Data 10−2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−5. Operating Instructions: Air Conditioning (P/N 900P7250302−103) 10−3. . . . . . . . . .

Figure 10−1. Air-conditioning System 10−4. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−2. Air Conditioner Control 10−6. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−6. Operating Instructions: Controllable Landing/Search Light 10−7. . . . . . . . . . . . . . . .

Table 10−3. Search Light Switch Functions 10−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−3. Collective Stick Switch Panel 10−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−4. Circuit Breakers − Baggage Compartment Mounted (Typical) 10−10. .

10−7. Operating Instructions: Rotorcraft Cargo Hook Kit 10−13. . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−5. VNE Placard 10−14. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−6. Weight and Balance Envelope 10−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−7. Cargo Hook IIDS Menu 10−17. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−8. Cargo Hook Installation 10−18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−8. Operating Instructions: Windscreen Wipers 10−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−9. Windscreen Wiper with Optional Windscreen

Washer Installation 10−21. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−10. Windscreen Wiper Control Switch 10−22. . . . . . . . . . . . . . . . . . . . . . . . . .

Servicing Materials − Windscreen Washer Fluid 10−23. . . . . . . . . . . . . . . . . . . . . . . . . .

10−9. Operating Instructions: Supplemental Fuel System 10−25. . . . . . . . . . . . . . . . . . . . . . .

Figure 10−11. Gauge, Switch and Indicator Light − Location Typical 10−27. . . . . . .

EXAMPLE I: Longitudinal CG Determination 10−28. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 10−4. Fuel Loading Table − Jet−A (6.8 LB/GAL) 10−29. . . . . . . . . . . . . . . . . . . . .

Table 10−5. Fuel Loading Table − Jet−B (6.5 LB/GAL) 10−29. . . . . . . . . . . . . . . . . . . . .

Figure 10−12. Fuel Station Diagram 10−30. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−13. Supplemental Fuel System 10−32. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Optional Equipment

FAA ApprovedReissue 1Original 10−ii

PARAGRAPH PAGE

10−10. Operating Instructions: Rescue Hoist 10−35. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−14. Center of Gravity Envelope for Hoist Operations Below 60 KIAS 10−36. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−15. Rescue Hoist Controls 10−40. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

EXAMPLE I: Lateral CG Determination − Enroute (above 60 KIAS) 10−42. . . . . . . . EXAMPLE II: Lateral CG Determination − Destination (below 60 KIAS) 10−42. . . . EXAMPLE III: Lateral CG Determination − With Hoist Load 10−43. . . . . . . . . . . . . . Figure 10−16. Allowable Rescue Hoist Loading Chart 10−44. . . . . . . . . . . . . . . . . . . . .

Figure 10−17. Rescue Hoist Installation 10−46. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 10−6. Servicing Materials 10−47. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−11. Operating Instructions: Removable CoPilot Controls 10−49. . . . . . . . . . . . . . . . . . .

Figure 10−18. Collective and Cyclic Placards 10−49. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−19. Removable Copilot Cyclic Control 10−51. . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−20. Removable Copilot Collective Control 10−52. . . . . . . . . . . . . . . . . . . . . . .

10−12. Operating Instructions: Airframe Fuel Filter 10−55. . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−21. Airframe Fuel Filter Installation and Block Diagram 10−58. . . . . . . .

10−13. Operating Instructions: SX−16 Night Sun with AftMount 10−59. . . . . . . . . . . . . . .

Figure 10−22. SX−16 Aft Mount Installation 10−61. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−23. SX−16 Searchlight Assembly 10−63. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−14. Operating Instructions: RDR−1400C Weather Radar 10−65. . . . . . . . . . . . . . . . . . .

Figure 10−24. RDR 1400C EFIS System Interface 10−67. . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−25. CP 113 10−69. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−15. Operating Instructions: LEO−II−A5 Observation System 10−71. . . . . . . . . . . . . . .

Figure 10−26. LEO−II−A5 Mounting 10−72. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−16. Operating Instructions: Annunciator panel 10−75. . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−27. Caution and Advisory Annunciators 10−75. . . . . . . . . . . . . . . . . . . . . . . .

10−17. Operating Instructions: Moving Map Navigation Systems 10−77. . . . . . . . . . . . . .

Figure 10−28. AVM4090 Display Controls 10−78. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 10−7. AVM4090 Display Control Functions 10−78. . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−29. AI−500 Monitor 10−79. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table 10−8. AI−500 Control Functions 10−79. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−18. Operating Instructions: W.E.S.T Battery Protection System 10−81. . . . . . . . . . . . .

Figure 10−30. Ensave02 Indicator/Button 10−81. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−19. Operating Instructions: SX−16 Night Sun with EPMS Mount and Laser Pointer 10−83. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−31. SX−16 EPMS Installation 10−84. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Optional Equipment

FAA ApprovedReissue 1Original 10−iii/( 10−iv blank)

PARAGRAPH PAGEFigure 10−32. W.E.S.T. SX−16 Control Panel 10−86. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

10−20. Operating Instructions: Smoke Detector 10−89. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CArgo Compartment Fire/Smoke 10−89. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−33. Smoke Detector and Press−To−Test Switch Location. 10−90. . . . . . . . .

10−21. Operating Instructions: Crew Door Modification withQuick Release Mechanism 10−93. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 10−34. Cockpit Door Attachment 10−95. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .


Optional Equipment




SECTION XOPTIONAL EQUIPMENT

10−1.GENERAL INFORMATION

This section provides general supplemental information on optional equipment forthe helicopter. The information includes a listing of usable optional equipment andcompatibility of combined equipment on the helicopter.

Supplemental data is prepared and included in this section whenever the installationof that equipment affects the FAA Approval Data for Limitations (Section II), Emer-gency and Malfunction Procedures (Section III), Normal Procedures (Section IV),and Performance Data (Section V).

The Flight Manual Supplemental Data is to be used in conjunction with the basicFlight Manual data and takes precedence over that data when the equipment isinstalled.

Be sure to include a review of the appropriate flight manualsupplemental data for type of optional equipment installed(including STC items) as a regular part of preflight planning.

10−2.LISTING − OPTIONAL EQUIPMENT

Table 10−1 lists MDHI optional equipment items available that require additionaloperatinig instructions. This table does not include non−MDHI STC items that maybe FAA approved for use. Other optional equipment items may be found in theRMM.

SPECIAL NOTE:

Items in the table marked with an asterisk (*) are optional equipment items thathave had their supplemental data incorporated into the main body of the flightmanual and are identified by the statement, ‘‘If installed’’.

CAUTION


MD900 (902 Configuration with PW 207E)Optional Equipment


Table 10−1. Optional Equipment MD900 HelicopterEquipment Publication No.

Air conditioner (P/N 900P7250302−101) CSP−902RFM207E−1 Section X

Search light CSP−902RFM207E−1 Section X

Cargo hook CSP−902RFM207E−1 Section X

Windscreen Wipers CSP−902RFM207E−1 Section X

Supplemental Fuel System CSP−902RFM207E−1 Section X

Rescue Hoist CSP−902RFM207E−1Section X

*Pitot heat CSP−902RFM207E−1

*Rotor brake CSP−902RFM207E−1

*Engine air particle separator filter CSP−902RFM207E−1

* Indicates data incorporated into the flight manual (Sections I thru IX where appropriate).

10−3.COMPATIBILITY − COMBINED OPTIONAL EQUIPMENT

Table 10−2. Optional Equipment Kit Compatibility − MD900 Helicopter

Compatibility: Blank = Yes; X = No

Optional Equipment

A. B. C. D. E. F. G. H. I. J. K. L. M. N.

A. Air conditioner

B. Search light

C. Engine air particle separator

D. Rotor brake

E. Pitot heat

F. Cargo hook

G. Windscreen Wipers

H. Supplemental Fuel System

I. Rescue Hoist

J. Removable Copilot controls

K. Airframe fuel filter

L. SX−16 Searchlight

M. RDR 1400C Weather radar

N. LEO−II−A5 Observation system

10−4.OPTIONAL EQUIPMENT PERFORMANCE DATA

SPECIAL NOTE:Optional equipment that affect IGE/OGE hover performance requireadditional hover performance charts. All Optional Equipment hoverperformance charts are located in Section V.

Optional EquipmentAir−Conditioning System



10−5.OPERATING INSTRUCTIONS: AIR CONDITIONING (P/N 900P7250302−103)

PART IGENERAL

The air−conditioning system circulates conditioned air throughout the cabin. A fiveposition rotary switch AC/VENT controls the vent fan and air−conditioning. COOLHIGH provides air−conditioning at a high setting. COOL LOW provides air−condi-tioning at a low setting, selected from the center console utility panel assembly.The air−conditioning system provides ventilation, temperature, and humidity con-trol. The air−conditioning system consists of:

Freon Compressor Assembly − Compresses the air conditioning system refriger-ant.Lines and Tubing − Routes refrigerant throughout the air conditioning system.Condenser − Heat exchanger for the condensing refrigerant.Receiver Dehydrator − Removes moisture from the air conditioning system refrig-erant.High Pressure Switch − Turns off the compressor in a high pressure conditionto prevent damage to air conditioning system.Low Pressure Switch − Activates or deactivates the Freon compressor assemblyin a low pressure condition to prevent damage to the air conditioning system.Thermal Expansion Valve − Regulates air conditioning system refrigerant injectedinto the evaporator.Evaporator − Heat exchanger that cools cabin air.Evaporator Fan − Induces airflow through evaporator.Three Way Valve Duct Assembly − Controls the flow of recirculated cabin airor ambient air to the air conditioning system.Three Way Valve Control Cable − Controls position of the three way valve.

The compressor is mounted on the gearcase of the right hand engine. The condensersare attached to the oil cooler blowers. The evaporator occupies the forward endof the upper cowling. The air conditioning system makes use of the ventilation sys-tem’s ducting to direct the cooled air to cabin and cockpit, but adds a manual valveto permit selection of fresh or recirculated air. The knob for this push/pull controlis on the rear cockpit wall above the pilot’s right shoulder. Other air conditionercontrols are located on the Utility panel.




F927−053

LOW PRESSURE SWITCH

WATER SEPERATOR

3−WAY VALVE

3−WAY VALVEDUCT ASSEMBLY

3−WAY VALVECONTROL CABLE

EVAPORATORFAN

EVAPORATOR

THERMOSTATICSWITCH AIR PLENUM

DUCT

LINES ANDTUBING

SIGHTGLASS

COMPRESSOR

HIGH PRESSUREGUAGE SWITCH

RECIEVERDEHYDRATOR

RH CONDENSER

Figure 10−1. Air-conditioning System




PART IILIMITATIONS

No change.

PART IIIEMERGENCY AND MALFUNCTION PROCEDURES

NOTE: An automatic cutoff procedure similar to that for the heat/defog system shutsdown the air conditioner in flight if either engine becomes inoperative to maintainthe best power output from the running engine.

LOSS OF COOLING

Indications: No cooling air with system ON

Conditions: Automatic system safety shutdown, or internal failure

Procedures:

• A/C control switch OFF

• Use fresh air vent system as required




PART IVNORMAL PROCEDURES

A five position rotary switch AC/VENT controls the vent fan and air−conditioning.Selecting the COOL HIGH position provides air−conditioning at a high setting;COOL LOW provides air−conditioning at a low setting, selected from the centerconsole utility panel assembly.

IPS

HEAT

AC/VENT

PITOTHEAT

HYDTEST

OFF

ON

SYS 1 OVRD

SYS 2

COOLLOW

COOLHIGH

VENTHIGH

VENTLOW

OFF

OFF

OFF

ON

ON

L VSCS RON

OFF

TEST

CAB

AC/VENT CONTROL

F92−138

Figure 10−2. Air Conditioner Control

PART VPERFORMANCE DATA

Ref. Section V for hover performance with air-conditioning ON.

Optional EquipmentControllable Landing/Search Light




10−6.OPERATING INSTRUCTIONS: CONTROLLABLE LANDING/SEARCH LIGHT

PART IGENERAL

The controllable search light may be located on the lower fuselage ahead of theforward landing gear crosstube and offset to the left of the centerline or installedon a mounting pod that allows the use of the searchlight when other equipmentis installed in front of the standard search light location.

In the stowed position, the search light is flush with the lower fuselage skin andfaces downward.

Illuminating the search light is accomplished through the search light power switch(SRCH) while positioning the search light is accomplished by operating the five−posi-tion search light control switch (Ref. Figure 10−3). The search light is availablewith an optional IR lamp.

PART IILIMITATIONS

No Change.


No Change.






Table 10−3. Search Light Switch Functions

SWITCH POSITION FUNCTION

SRCH LT

OFF

Switches search light ON.

Switches search light OFF.

IR Switches IR lamp ON (if installed).

Search LightControlSwitch

EXT

RET

L

R

Press and hold switch to extend search light.

Press and hold switch to retract search light.

Press and hold switch to rotate search light to theleft.

Press and hold switch to rotate search light to theright.

Preflight Checks − Electrical power OFF:

� Search light CHECK FOR BROKEN COVER,DAMAGE TO MECHANICALASSEMBLY OR BURNED OUTBULB.

� Baggage compartment mounted SRCH LGT andHVR LGT circuit breakers (early configuration)

IN (REF. FIGURE 10−4)

� Baggage compartment mounted SRCH LT andSEARCH LT CONT circuit breakers (current con-figuration)

IN (REF. FIGURE 10−4)





Preflight Checks − Electrical power ON:

NOTE: The following operational checks shall be performed with an external powersource.

� Electrical master panel

� � POWER switch BAT/EXT

� Collective stick switch panel

� � SRCH switch to LT (Ref. Figure 10−3) CHECK SEARCH LIGHT ON

� � Use search light switch to rotate light left (L)and right (R)

CHECK OPERATION

� � SRCH switch to OFF SEARCH LIGHT OFF

F927−033

SEARCH LIGHT

SEARCH LIGHTPOWER SWITCH

CONTROL SWITCH

Figure 10−3. Collective Stick Switch Panel





F92−141A

PITOTHEAT 2

LEFT GENERATOR BUS

AUDIOPNL 2

CKPTUTL

CABUTL

EVAP VENTEVAP

COMP

ATTGYRO 2

CPLTCLOCK

CNDSRFAN 2

EADIL

EHSIL

L W/SWIPER

AHRS1 AUX

LH DCFDR

IIDS TRAKSTB

HYDTEST

AVFAN

IPS

HOIST CUT

HOISTPWR

ATTGYRO1

PILOTCLOCK

CNDSRFAN 1

ELT R W/SWIIPER

AHRS2 AUX

RH DCFDR

FD SYN FLT DIR MODE SEL INVTRLEFT ESS BUS

LEFT AVIONICS BUS

RIGHT GENERATOR BUS

ADF2 RADARRT

RADARIND

MKRBCN

RADALT

PAPWR

COM 3 XPNDR2

DIRGYRO 2

NAV 3

MVGMAP

LIGHTING

RIGHT AVIONICS BUS

L R L R

L R L R

BST PUMP EECRH FUEL

LOW

DETENT IGNTR

CNSL POSN STROBE AREA

AHRS2 PRI

AVMSTR

AUXFUEL

FIREHRD

SMOKEDET

ENCALT

SRCHLGT

HVRLGT

NACA LH FUELFUEL

CABAUD

5 VDIM

NSUNCONT

NSUNPWR

CARGOHOOK

L FLDEXCIT

R FLDEXCIT


ADF1 FMCTRL

FM1RT

FM2RT

FM3RT

DME STORMSCOPE

CAMERA NAV 1 RMI

BATTERY BUS

20

SEARCHLT CONT

SEARCHLT

CURRENT CONFIGURATION

CONFIGURATIONEARLY

Figure 10−4. Circuit Breakers − Baggage Compartment Mounted (Typical)






No change.

PART VIWEIGHT AND BALANCE DATA

No change.

PART VIISYSTEM DESCRIPTION

The search light is controlled by the three position SRCH toggle switch. This switchconnects battery bus power to the search light. Later configurations are poweredthrough the left generator bus.

Movement of the search light is accomplished by actuating the search light controlswitch located on the collective stick switch panel.

Maximum light extension is 120° from stowed.

If the search light is rotated 90° either side of center and with an extended segmentof 0° to 60°, an interlock switch automatically deenergizes the lamp while positioningthe light is still possible.

PART VIIIHANDLING SERVICING AND MAINTENANCE

No change.

PART IXADDITIONAL OPERATIONS AND PERFORMANCE DATA

No change.

Optional EquipmentRotorcraft Cargo Hook Kit




10−7.OPERATING INSTRUCTIONS: ROTORCRAFT CARGO HOOK KIT

PART IGENERAL

The cargo hook is an option that permits the helicopter to carry a jettisonable externalload of up to 3,000 pounds. The hook is suspended by a bridle of four cables thatattach to the landing gear saddle fittings, and join at the cargo load cell link tosupport the hook.

The pilot’s controls for the hook consist of an electric release push−button on thetop of the cyclic grip and a manual/emergency cargo hook release mechanism.

Quick disconnect pins at the four attachment points for the bridle allow the flightcrew to install or remove the hook assembly. Quick disconnects for the electric andmechanical release cables are located on the bottom of the fuselage near the forwardcross tube.

When the kit is installed, an owner or operator holding a valid Rotorcraft ExternalLoad Operator Certificate may utilize the helicopter for transportation of externalcargo when operated by a qualified pilot. OPERATIONS WITH CARGO ON THEHOOK SHALL BE CONDUCTED IN ACCORDANCE WITH APPLICABLE POR-TIONS OF FEDERAL AVIATION REGULATIONS PART 133.

Information provided in this supplement is presented with the intent of furnishingimportant data that can be used in the Rotorcraft Load Combination Flight Manual.The Combination Flight Manual, which is required by FAR Part 133, will be preparedby the applicant to obtain the rotorcraft External Load Operator Certificate.





PART IILIMITATIONS

Weight Limitations:

Maximum weight allowed on the landing gear is 6500 pounds.Weight in excess of 6500 pounds and up to 6900 pounds mustbe external and jettisonable.

Maximum Rotorcraft − Load Combination operating gross weight is 6900 pounds.

Center of Gravity Limitations:

See Figure 10−6.

Cargo Hook Limitations:

Maximum weight on the hook is 3000 LBS unless placarded otherwise (Ref.Figure 10−8).

Airspeed Limitations:

With no load on hook, maximum VNE is 90 KIAS.

With load on hook, maximum VNE is 100 KIAS (Ref. Figure 10−5).

NOTE: Use caution as size and shape of load, and load attaching cable size and lengthmay affect flight characteristics. Satisfactory flight characteristics have beendemonstrated with a compact load.

Placards: Placard located on instrument panel.

F92−142

20000

15000

10000

5000

0

40 50 60 70 80 90 100 110

DE

NS

ITY

ALT

ITU

DE

− F

EE

T


VNE WITH LOAD ON THE HOOK

VNE WITH NO LOAD ON THE HOOK IS 90 KIAS

Figure 10−5. VNE Placard

CAUTION





CG LIMITS >6250LBS

194 196 198 200 202 204 206

7000

6500

6000

5500

5000

4500

4000

3500

3000

FUSELAGE STATION (IN.)

WE

IGH

T −

PO

UN

DS

F92−143A

208

NORMAL CG LIMITS

CARGO HOOKCG LIMITS >6250LBS

7000

6500

6000

5500

5000

4500

4000

3500

3000

LATERAL CG STATION (IN)

WE

IGH

T −

PO

UN

DS

−5 −4 −3 −2 −1 0 1 2 3 4 5

NORMAL CG LIMITS

CARGO HOOK

LATERAL CGENVELOPE

LONGITUDINAL CGENVELOPE

Figure 10−6. Weight and Balance Envelope






The presence of an external load may further complicate procedures following anemergency or malfunction. Release of loads attached through the cargo hook shouldbe considered consistent with safety of flight factors.

Emergency Release:

Actuate the mechanical release handle, mounted on the cyclic stick, to releasecargo in the event of an electrical failure.


Preflight Checks (Ref. Figure 10−8):

Verify security of cargo hook bridle attach points.

Visually inspect hardware for damage or indications of possible fatigue.

Check for fraying, wear or any other form of damage to the cable bridle assembly.

Inspect electrical release, and load indicating wire harness and connectors for gener-al condition and security.

Examine manual release cable housing for nicks, cuts, kinks or general damagethat might restrict movement of cable within housing.

Inspect manual release connector for general condition and security.

Ensure a service loop is present in the manual release cable at cargo hook.

Inspect hook for general condition.

Cargo Hook Operational Checks:

NOTE: Functional checks of the cargo hook require an external power source forelectrical power or an operating engine.

Ensure that the CRGO HOOK circuit breaker (left generator bus) is IN.

NOTE: Refer to Chapter 25−55−00 in the RMM for special functional checks requiredfollowing the initial installation of the cargo hook kit or following replacement ofthe manual release cable.





With the load beam in its locked position, apply pressure to simulate a load onthe beam and functionally check the three methods of cargo hook release:

Mechanical release lever on the right side of hook

Manual cargo hook release handle on cyclic

Electric cargo hook release switch on cyclic

NOTE: The TARE weight should be reset each time following aircraft shutdown andrestart (Ref. Figure 10−7).

Operating Procedures:

Use care to avoid passing load attaching cables over landing gearskid tube when attaching load to hook with helicopter on theground.

Apply collective smoothly when lifting cargo.

With the hook weight suspended, and the selection made on the IIDS panelmenu for HOOK WT (Ref. Figure 10−7), the load indication should read HOOKWT. xxxx LBS on the alphanumeric display.

Ensure that there is adequate clearance between the sling load and any obstaclesalong the takeoff flightpath.

Activate cargo release switch on cyclic stick to release cargo.

Check CARGO HOOK OPEN advisory on IIDS alphanumeric display.

NOTE: Ground support personnel should manually assure positive reset of the cargohook after use of mechanical release, prior to further cargo pickups.

Instruct ground crew to ensure that the helicopter has been electrically groundedprior to attaching cargo to drain charges of static electricity that may build upin flight.

The cargo hook extends 18 inches below the landing gear whilehovering. Ensure that there is adequate clearance between thecargo hook and any obstacles along the flightpath.

SET CALIB CODE

<XXXX>

HOOK WT 2456 LBS ZERO WEIGHT DISP

2456 LBS



�ENT" SELECTS DIGITS TO BE EDITED (LEFT TO RIGHT),

AND �� KEYS INCREMENT/DEINCREMENT DIGIT VALUE,

�REC" KEY CHANGES CODE TO SELECTED VALUE,

PRESSING �ENT" FOR MORE THAN 2 SECONDS

TAKES A TARE READING AND ZEROS DISPLAY

F92−144

Figure 10−7. Cargo Hook IIDS Menu

CAUTION

CAUTION





LH AFT LANDINGGEAR FITTING

RH AFT LANDINGGEAR FITTING

RH FWD LANDINGGEAR FITTING

LH FWD LANDINGGEAR FITTING

CYCLIC STICK

MANUAL CARGO HOOKRELEASE

F92−145A

CARGO HOOK

LOAD INDICATORELECTRICAL CONNECTOR

MANUAL RELEASECABLE CONNECTION

ELECTRICAL RELEASECONNECTOR

AFT SADDLE CLAMP

FWD SADDLE CLAMP

QUICKRELEASE PIN

(SEE NOTE)

PIN LINK

CABLE

LOAD BEAM

QUICK RELEASE PIN(SEE NOTE)

FORWARDPIN LINK

CABLE

ELECTRIC CARGOHOOK RELEASE

CYCLIC GRIP ROTATED

MECHANICALRELEASE LEVER

SERVICE LOOP

NOTE: ENSURE QUICK RELEASEPIN HEAD FACES ‘‘UP’’AFTER INSTALLATION

LINK ASSEMBLY

LINK ASSEMBLY

CARGO HOOK PLACARD

MAX WORKINGLOAD 2200 LB

Figure 10−8. Cargo Hook Installation






Hover Ceiling:

Use the OGE hover ceiling charts: Refer to Section V for Hover Ceiling Data.


Cargo Hook Longitudinal CG: 203.0 In.

Cargo Hook Assembly Weight: 26.12 lbs.

The following table of Cargo Hook Loads may be used by the operator to assistin determining the helicopter center of gravity.

Cargo Weight(lb)

Moment/100(in.−lb)

Cargo Weight(lb)

Moment/100(in.−lb)

100 20300 1600 324800

200 40600 1700 345100

300 60900 1800 365400

400 81200 1900 385700

500 101500 2000 406000

600 121800 2100 426300

700 142100 2200 446600

800 162400 2300 466900

900 182700 2400 487200

1000 203000 2500 507500

1100 223300 2600 527800

1200 243600 2700 548100

1300 263900 2800 568400

1400 284200 2900 588700

1500 304500 3000 609000






Cargo Hook Installation (Ref. Figure 10−8):

Align cargo hook cable attaching hardware with landing gear saddle clamp assem-blies.

Install FWD link assemblies into FWD saddle clamps.

Install FWD pin links into link assemblies and quick release pins into FWDpin links.

Connect cargo hook electrical connector, load indicator electrical connector andmechanical release control cable connector.

Repeat procedure for aft link assembly attachment.

Perform cargo hook preflight and operational checks.

Cargo Hook Removal (Ref. Figure 10−8):

Remove quick release pins from Aft pin links and remove cable assembly fromaft saddle clamps.

Disconnect cargo hook electrical connector, load indicator electrical connectorand manual release control cable connector.

Remove pin links attaching cargo hook cables and cargo hook to FWD landinggear saddle clamp assemblies.

Remove cargo hook and bridle assembly.

Optional EquipmentWindscreen Wipers




10−21

10−8. OPERATING INSTRUCTIONS: WINDSCREEN WIPERS

PART IGENERAL

The windscreen wipers provide the pilot a means to clear the windscreens of rainor snow. The windscreen washers (if installed) provide pressurized washer fluid to the wind-screen through spray nozzles. The washer pump and reservoir are located in thebattery compartment.

There are no changes to limitations, emergency procedures, or performance datawith the installation of the windscreen wipers or windscreen washers.

F92−170A

WASHER PUMP

WASHERRESERVOIR

WINDSCREENWIPERS

Figure 10−9. Windscreen Wiper with Optional Windscreen Washer Installation





10−22


Windscreen wipers:

Use the windscreen wipers whenever it is necessary to clear the windscreensof rain or snow.

Do not use the windscreen wipers on a dry windscreen.

The wipers have a panel mounted control switch (Ref. Figure 10−10). The rotary-switch has four positions:PARK, OFF, LOW, and HIGH.The PARK position is a momentary position and is used to stow the wipers whennot in use.The OFF position turns the wipers off.The LOW and HIGH positions refer to wiper speed. Select the speed appropriatefor weather conditions.

The three position toggle switch HIGH (HI) and LOW (LO) positions functionas above. The OFF position parks and turns the wipers off.

Windscreen washer (if installed):

Preflight Check

Check washer reservoir fluid level.

On dry windscreen

Press and hold the WASHER button for two to three seconds before turningthe WINDSHIELD WIPERS switch to LOW. Turn off wipers while windscreenis still wet.

During wiper operation

Press and hold the WASHER button for two to three seconds or as neededto clear the windscreen.

Cold weather operation

Use 50 percent by volume isopropyl alcohol mixed with distilled or deionizedwater when temperatures are at or below 0°C.

F927−093B

ROTARY CONTROL SWITCH

PARKOFF

LOW

DO NOT OPERATE WIPERS

WASHER

ON DRY WINDSCREEN

WINDSHIELDWIPERS

HIGH

WASHER CONTROL SWITCH(IF INSTALLED)

OFF

LOW

DO NOT OPERATE WIPERSON DRY WINDSCREEN

WINDSHIELDWIPERS

HIGH

3−POSITION TOGGLE SWITCH

WIPERS

HI

LO

OFF

DO NOT OPERATE WIPERSON DRY WINDSCREEN

3−POSITION LOCKINGTOGGLE SWITCH

Figure 10−10. Windscreen Wiper Control Switch





10−23/(10−24 blank)


ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)

Washer reservoir full − water onlyWasher reservoir full − water alcohol mixture

4.84.3

82.782.7

394356

PART VIIIHANDLING, SERVICING AND MAINTENANCE

Servicing Materials − Windscreen Washer Fluid


Washer reservoir − Total Capacity approximately 2 US quarts.

None Distilled or deionized water for opera-tions above freezing and 50 percent byvolume mixture of isopropyl alcoholand distilled or deionized water for op-erations below freezing.

None

Optional EquipmentSupplemental Fuel System



10−25

10−9. OPERATING INSTRUCTIONS: SUPPLEMENTAL FUEL SYSTEM

PART IGENERAL

The MD900 supplemental fuel system option adds a transfer type auxiliary fueltank located below the baggage compartment floor. Refer to Part VII for systemdescription.

PART IILIMITATIONS

Placards:

SUPPLEMENTAL FUEL SYSTEMUSE MAIN FUEL DOWN TO

700 LBS BEFORE SELECTINGAUX FUEL TRANSFER

LOCATED BY AUXILIARY FUEL GAUGE.

NOTE: LOCATION MAY VARY.

LOCATED ABOVE FUEL FILLERF92−172A




10−26


� Preflight checks:

� � Fuel cap SECURED� Prestart cockpit check:

� � Fuel transfer switch OFF� Inflight operation:

� � Fuel transfer switch ON; VERIFY FUEL TRANSFER LIGHT‘‘ON’’

NOTE: Fuel transfer should be begun when the fuel level in the main tank is between700 and 300 LBS.

� � Main fuel tank quantity begins toincrease and auxiliary fuel quantitybegins to decrease.

CHECK

NOTE: Fuel Transfer:Fuel transfer time is approximately 20 minutes (22 minutes if second check valveinstalled) with a full auxiliary fuel tank while in normal cruise. Transferring fuel to the main tank may be accomplished once main tankindicated fuel quantity is at or less than approximately 500 LB in normal groundattitude or approximately 700 LB in normal cruise attitude.Fuel transfer rate is approximately 600 LB/HR (540 LB/HR with second checkvalve) in normal cruise and approximately 400 LB/HR in normal ground attitude.

Starting Fuel Transfer below 300 LBS:With engines at MCP, the auxiliary fuel transfer rate may not keepup with the engine fuel consumption rate.

Starting fuel transfer below 300 LBS following a boost pump failure(boost pumps OFF, Ref. Section III) may result in early right enginefuel starvation (fuel transfers from the auxiliary fuel tank into theleft side of the main fuel tank).

� � Fuel transfer switch OFF WHEN TRANSFER IS COMPLETE

NOTE: The auxiliary fuel quantity gauge has been found to be inaccurate (indicateshigh) during hover operations.

� Engine/aircraft shutdown

� � Fuel transfer switch OFF; VERIFY FUEL TRANSFER LIGHT‘‘OFF’’

CAUTION




10−27

F92−173B

ON

OFF

E F

AUX FUEL

50 100 150

FUEL TRANSFER INDICATOR LIGHT

FUEL TRANSFER SWITCH

AUXILIARY FUEL QUANTITY GAUGE

GAUGE, SWITCH AND INDICATOR LIGHTS

SUPPLEMENTAL FUEL SYSTEMUSE MAIN FUEL DOWN TO

700 LBS BEFORE SELECTINGAUX FUEL TRANSFER.

FUELXFER

Figure 10−11. Gauge, Switch and Indicator Light − Location Typical




10−28


Weight and balance characteristics:

The lateral CG of the auxiliary fuel tank is at station −2.0.

Calculate CG as shown in the example below.

EXAMPLE I: Longitudinal CG Determination

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)

Basic Weight 3512.4 738045

Pilot 185.0 130.70 24180

Copilot/Passenger 185.0 130.70 24180

Passenger − Rear Facing R/H 175.0 173.0 30275

Passenger − Rear Facing L/H 175.0 173.0 30275

Passenger − FWD Facing R/H 175.0 213.0 30275

Passenger − FWD Facing L/H 175.0 213.0 30275

1. Zero Fuel Weight 4582.4 201.1 30275

2. Add: Fuel − Main Tank Only (Jet−A)Gross Weight:

1025.05607.4

191.2199.3

1959801117484

3. Add: Fuel − Auxiliary Tank OnlyGross Weight:

200.04782.4

244.8202.9

48960970464

4. Add: Fuel − Both TanksGross Weight:

1225.05807.4 200.9 1166444




10−29

Table 10−4. Fuel Loading Table − Jet−A (6.8 LB/GAL)

FUEL WEIGHT(LB)

LONGITUDINALSTATION

LONGITUDINALMOMENT

LATERALSTATION

LATERALMOMENT

20 239.4 4789 −2.0 −40

40 240.6 9625 −2.0 −80

60 241.6 14494 −2.0 −120

80 242.3 19387 −2.0 −140

100 242.9 24294 −2.0 −200

120 243.4 29210 −2.0 −240

140 243.8 34133 −2.0 −280

160 244.1 39062 −2.0 −320

180 244.4 44000 −2.0 −360

200 244.8 48951 −2.0 −400

Table 10−5. Fuel Loading Table − Jet−B (6.5 LB/GAL)

FUEL WEIGHT(LB)

LONGITUDINALSTATION

LONGITUDINALMOMENT

LATERALSTATION

LATERALMOMENT

20 239.5 4790 −2.0 −40

40 240.7 9629 −2.0 −80

60 241.7 14501 −2.0 −120

80 242.5 19397 −2.0 −140

100 243.1 24306 −2.0 −200

120 243.5 29224 −2.0 −240

140 243.9 34149 −2.0 −280

160 244.3 39080 −2.0 −320

180 244.6 44022 −2.0 −360

200 244.9 48982 −2.0 −400




10−30

210200190180170160150140130120110100908070605040302010

0

FU

EL

WE

IGH

T −

LB

JET − B (6.5 LB/GAL)

FUSELAGE STATION − INCHES238.0 238.5 239.0 239.5 240.0 240.5 241.0 241.5 242.0 242.5 243.0 243.5 244.0 244.5 245.0

210200190180170160150140130120110100908070605040302010

0

FU

EL

WE

IGH

T −

LB

JET − A (6.8 LB/GAL)

FUSELAGE STATION − INCHES

238.0 238.5 239.0 239.5 240.0 240.5 241.0 241.5 242.0 242.5 243.0 243.5 244.0 244.5 245.0

F92−174

Figure 10−12. Fuel Station Diagram




10−31


The MD900 supplemental fuel system option adds a transfer type auxiliary fueltank with a usable capacity of approximately 29.4 US gallons (200 LB, Jet−A) under-neath the baggage compartment floor. The tank is filled through a gravity fill porton the right side of the aircraft. Transfer into the main tank is performed usinga fuel transfer pump mounted in the auxiliary fuel tank. Overfilling the main tankis prevented by use of a float−type level control valve mounted in the main tank.This level control valve prevents transfer into the main tank until the fuel remainingin the main tank is less than approximately 500 LB in normal ground attitudeor approximately 700 LB in normal cruise attitude. The level control valve willshut off transfer into the main tank if the fuel in the main tank increases to approxi-mately 755 LB in normal ground attitude or approximately 832 LB in normal cruiseattitude. An inline check valve is installed in the fuel transfer line in the mainfuel tank to prevent backflow of fuel from the main tank into the auxiliary tank.A second check valve may be installed in the auxiliary fuel tank transfer line thatprevents gravity transfer from the auxiliary tank into the main tank in high−speedcruise flight. The auxiliary tank vent is teed into the existing main tank aft venttubing.

The installation includes a cockpit mounted fuel quantity gauge (AUX FUEL) forthe auxiliary tank, a fuel transfer pump switch, and a fuel transfer indicator lightand on some installations, an additional grounding jack is located above the auxiliaryfuel tank filler.

Electrical power is supplied from the battery bus through the ‘‘AUX FUEL’’ 5 AMPcircuit breaker. A separate 1 AMP ‘‘AUX FUEL XMIT’’ circuit breaker providespower for fuel quantity indicating. These circuit breakers are located on the baggagecompartment circuit breaker panel under ‘‘BATTERY BUS’’.




10−32

AUXILIARY TANK FILLER NECK

MAIN FUEL TANK (REF.)

FUEL VENT LINE

FUEL LEVEL CONTROL VALVE

FUEL TRANSFER LINE

AUXILIARY FUEL TANK

FUEL TRANSFER LINE

FUEL QUANTITY TRANSMITTER

FUEL VENT ROLL OVER VALVE

GRAVITY FILL PORT

AFT RH VENT LINE

FLAME ARRESTOR

AFT VENT FAIRING

AUX FUEL PORTVENT ROLLOVER VALVE

CABIN FLOOR (REF)

MAIN FUEL TANK(REF)

LEVEL CONTROL VALVE

STA 230.5 BULKHEAD

CHECK VALVE

FUEL TRANSFER LINE

BAGGAGE COMPARTMENT FLOOR(REF.)

VENT/ROLLOVER VALVE

CHECK VALVE

AUXILIARY FUEL TANK

FUEL QUANTITY XMITTER FLOAT

TRANSFER PUMP

FUEL TANK DRAIN PLUG

FUEL TRANSFER LINE

FUEL TRANSFER LINE VENT LINE

(AUXILIARY FUEL TANK ROTATED 90° CW FOR CLARITY)

SUPPLEMENTAL FUEL SYSTEM SCHEMATIC

SUPPLEMENTAL FUEL SYSTEM INSTALLATION

F92−175B

NEW CHECK VALVE(IF INSTALLED)

Figure 10−13. Supplemental Fuel System




10−33/(10−34 blank)


Fuel additives:

Anti−icing additives, if required, must be added to the auxiliary fuel tank duringrefueling.

Optional EquipmentRescue Hoist




10−10.OPERATING INSTRUCTIONS: RESCUE HOIST

PART IGENERAL

The rescue hoist system provides a means for lowering and raising personnel orcargo from an airborne helicopter. It is capable of being operated from the passengercabin by a qualified crewmember or from the pilot’s station.

PART IILIMITATIONS

Type of operation:

Hoist operations shall be conducted under appropriate airworthiness and/or oper-ating rules for external loads.

Minimum flight crew:

Pilot, when conducting operations with hoist stowed.

Pilot and hoist operator, when conducting hoist operations.

NOTE: Hoist operator must wear appropriate safety gear, safety harness, and havevoice communications with the pilot during hoist operations.

Weight and balance:

Maximum lateral CG limit at 60 KIAS or less:+9.0 IN at 5550 LB gross weight; +7.5 IN at 6500 LB gross weight.

At airspeeds above 60 KIAS, normal CG limits apply.

With hoist installed, lateral C.G. may be exceeded with fuelconsumption. Flight planning should include a minimum fuel lateralC.G. check.

CAUTION





Airspeed limitations:

Observe airspeed limitations in Section II with hoist installed and doors closed.

Observe VNE for doors open/removed flight in Section II.

VNE while conducting hoist operations is 60 KIAS.

Hoist limitations:

Maximum load on hoist is 600 LB.

Maximum permissible cable deflection is 15° with respect to the aircraft verticalaxis.

During normal flight operations and airspeeds above 60 KIAS, the cable/hookmust be in the fully raised position.

Center of gravity limitations:

Size, weight, shape of load and cable length may affect flightcharacteristics.

3000

3500

4000

4500

5000

5500

6000

6500

−3 −2 −1 0 1 2 3 4 5 6 7 8 9 10LATERAL C.G. STATION (IN.)

WE

IGH

T −

PO

UN

DS

F927−134

MAXIMUM LATERAL CGLIMIT AT 60 KIAS OR LESS:+9.0 IN AT 5550 LB GROSSWEIGHT; +7.5 IN AT 6500 LBGROSS WEIGHT.

Figure 10−14. Center of Gravity Envelope for Hoist Operations Below 60 KIAS

CAUTION






CABLE CUTTING:

Procedures:

Pilot: Activate the CABLE CUT switch on collective to jettison load in theevent of an emergency.

Hoist operator: Use provided cable cutters.

GENERATOR FAILURE:

NOTE: Hoist operations can require up to 125 amps of electrical power (63% load fromone generator).

Procedures:Monitor operating generator load and turn off unnecessary electricalequipment if required to maintain generator load within limits. Allowinga GENERATOR HIGH LOAD condition to exist will result in theoperating generator going off line.

ADVISORY INDICATIONS:

Indications: Green or yellow indicator light (located on control pendant) − steady greenor yellow

Indications:

Conditions: Motor overtemperature.

NOTE: The light will remain on until the motor has cooled or electrical power to thecontroller is switched off.

Procedures: Complete hoist operation in progress.

Prolonged operation of hoist with motor overtemperature lightilluminated will result in damaged or a ‘‘burned out’’ motor.

Indications: Flashing green light and a reduction of hoist speed.

Conditions: Hoist load above 250 LB with load mode select switch set to 250.

Procedures: Reduce hoist load or place load mode select switch to 600.

CAUTION






Preflight checks:

To lower the work platforms/steps, remove the quick release pinon the hoist strut and move the strut aside. The quick release pinmust be reinstalled before any load is placed on the hoist.

NOTE: External power is required for functional checks.� Rescue hoist assembly CHECK − FOR OIL

LEAKS AND GENERALCONDITION

� Hoist fairing CONDITION ANDSECURITY

� Electrical connections CHECK

� Hook assembly − freeness of swivel and latch CHECK

� Hoist support tube CHECK MOUNTING

� Hoist strut CHECK MOUNTING ANDQUICK RELEASE PIN

� Pendant control − electrical connection CHECK

� HOIST PWR and HOIST CUT circuit breakers IN

� Pilot’s hoist control panel CHECK SWITCH OFF

� Electrical Master Panel

� � Power switch BAT/EXT

� Pilot’s Hoist Control

� � Hoist arming switch ON

� � Hoist armed light − on CHECK

� � Payout displays CHECK

� � Hoist operational check (pilot and operator) CHECK

Do not restrict cable payout during this check. Fouling of the cableon the drum will result if this precaution is not followed.

� � Reel out approximately 25 feet (8 meters) or more ofcable by using both the pendant and the pilot payoutcontrols. Do not exceed 15° cable deflection

OPERATE HOIST

CAUTION

CAUTION





NOTE: The cable should be reeled out onto a smooth, clean surface or payed out intoa drum. Exercise care to prevent kinking of the cable.

� � Reel in cable by using both the pendant and the pilotpayout controls and verify hoist stops when hookreaches upper limit without excess tension on cable.Verify that pilot’s pay out switch overrides hoist oper-ator’s pendant control

OPERATE HOIST

NOTE: It is important that the cable be reeled in with an even pull under a drag load of10 to 20 LB so that it does not wrap loosely on the drum. A drag load must beapplied using a gloved hand or clean heavy cloth on the cable to achieve tight,even layers on the drum.

� � Hoist arming switch OFF

� Electrical Master Panel:

� � Power switch OFF

Hoist operation:

Hoist operator must wear appropriate safety gear, safety harness,and have voice communications with the pilot during hoistoperations.

NOTE: Operation of the pilot’s payout switch overrides the hoist operator.

� Hoist arming switch ON

� Stabilize the aircraft in a hover over area ESTABLISH

� Cabin door (if closed) OPEN

� Hoist operator select load mode 250 OR 600 LB

� Payout control switch DOWN

NOTE: If possible, ensure that the helicopter has been electrically grounded prior toattaching cargo to drain static electricity that may build up in flight.

� Payout control switch UP

� Maintain hover until load is inside passenger cabin unless safety or operationalconditions dictate otherwise.

NOTE: Certain combinations of weight and cable length may induce a noticeablelateral oscillation. Should a lateral oscillation occur, raise or lower the load toalleviate this condition.

WARNING





ÔÔÔÔÔÔÔÔÔÔ

OFF

UP

DNHOIST

MOTOR

WRN

250

600

LB

FEET

106

CABLE PAYOUTFEET

106

PILOT’S HOIST CONTROL

INDICATORLIGHT

LOAD MODE SELECT

CABLE PAYOUT DISPLAY

ICS CONTROL SWITCH

HOIST OPERATOR’S CONTROLPENDANT ASSEMBLY

ON

OFFHOIST

PAYOUTDIRECTION/SPEED

CONTROL

HOIST ARMING SWITCH

COLLECTIVE CONTROLMODULE (REF)

DN

UP

PILOT’S PAYOUTCONTROL SWITCH

HOISTCUT

HOISTPWR

BAGGAGE COMPARTMENT MOUNTEDCIRCUIT BREAKER PANEL

RESCUE HOISTCIRCUIT BREAKERS

HOIST ARMED LIGHT F927−058B

EMERGENCY CABLE CUT

HOIST

CABLE CUT

T/O

TIMER

ÔÔÔÔÔÔÔÔÔÔ

OFF

UP

DNHOIST

MOTOR

WRN

250

600

LB

METERS

25

25

CABLE PAYOUTMETERS

HOIST

PWR

HOIST POWER SWITCH/ARMEDINDICATOR

LATEST CONFIGURATION

NOTE: LOCATIONS OF PILOT’S HOIST POWER SWITCH/ARMED LIGHT AND HOISTPOWER SWITCH/ARMED INDICATOR VARY WITH INSTALLED OPTIONS.

CABLE

CUT

LATESTCONFIGURATION

EMERGENCY CABLE CUTPUSHBUTTON SWITCH

Figure 10−15. Rescue Hoist Controls






Reduce hover gross weight capability 70 LB when hovering withrescue hoist installed.

Refer to Section V for hover performance data.


Maximum operating and hoist load weights:

Maximum gross weight for hoist operations is 6500 LB including hoist load.

Maximum load on the hoist is 600 LB. This is a structural limit and does notassure loading within approved limits. Maximum allowable hoist load changeswith gross weight and aircraft CG. Refer to Figure 10−16 to determine maximumallowable hoist load.

ITEM WEIGHTSTATION (ARM) MOMENT

Lateral Longitudinal Lateral Longitudinal

Hoist installation 136.8 55.60 199.1 7611 27231

Hoist Load −−−−− 59.25 199.1 −−−−− −−−−−

Hoist lateral CG determination:

The following examples show a minimum crew of pilot and hoist operator. Noticethat in Example I, the helicopter is enroute (above 60 KIAS) and the hoist operatoris stationed in the left rear facing seat, thereby maintaining the lateral CG limitof �2 IN.

In Example II, the helicopter is at the destination (below 60 KIAS) and the hoistoperator moves to the right of the aircraft cabin and stands on step.

Example III shows CG with a load on the hoist.

CAUTION





EXAMPLE I: Lateral CG Determination − Enroute (above 60 KIAS)

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)


Hoist Installation 136.8 55.60 7611

Pilot 200 15.85 3170

Hoist Operator (L/H seat) 200 −19.00 −3800

Fuel 700 0.00 0

Gross Weight 4509.6 8446

Calculation of Lateral CG:

CG at Gross Weight:


8446= 1.90

Gross Weight 4509.6

EXAMPLE II: Lateral CG Determination − Destination (below 60 KIAS)

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)


Hoist Installation 136.8 55.60 7611

Pilot 200 15.85 3170

Hoist Operator (R/H step) 200 35.00 7000

Fuel 400 0.00 0

Gross Weight 4209.6 19246


CG at Gross Weight:


19246= 4.60

Gross Weight 4209.6





EXAMPLE III: Lateral CG Determination − With Hoist Load

ITEM WEIGHT(LB)

STATION(ARM)

MOMENT(IN−LB)

Basic Weight 3272.8 1465.0

Hoist Installation 136.8 55.60 7611.0

Pilot 200.0 15.85 3170.0

Hoist Operator (R/H step) 200.0 35.00 7000.0

Hoist load 250.0 59.25 14812.5

Fuel 400.0 0.00 0

Gross Weight 4409.6 34058.5


CG at Gross Weight:


34058.5= 7.72

Gross Weight 4409.6





F927−133

4700

6300620061006000590058005700560055005400530052005100500049004800

46004500440043004200410040003900

MAXIMUMHOIST LOAD

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0

600 LB500 LB

300 LB

HELICOPTER LATERAL STGATION CG (IN) − WITHOUT HOIST LOAD

HE

LIC

OP

TE

R G

RO

SS

WE

IGH

T (

LB)

− W

ITH

OU

T H

OIS

T L

OA

D

FOR USE BELOW60 KIAS ONLY

64006500

200 LB100 LB

400 LB

Figure 10−16. Allowable Rescue Hoist Loading ChartUse of chart:

Use Figure 10−16 to determine the maximum hoist load for this operation.

Example:

Known:

From EXAMPLE II: lateral CG = 4.6 inchesgross weight = approximately 4210 LB.

Enter chart at the ‘‘Helicopter Gross Weight Without Hoist Load’’ scale at 4210pounds and proceed horizontally to intersect with a line drawn vertically fromthe ‘‘Helicopter Lateral Station Without Hoist Load’’ scale at 4.6 IN. Where thetwo lines intersect is the allowable hoist load. For this example the allowablehoist load is approximately 369 pounds.






The hoist assembly consists of a cable drum that holds 245 feet (75.7 meters) of3/16" in. (5 millimeter) spin resistant cable, a fail safe load brake, 28 VDC electricmotor, limit switches coupled to the cable drum to control fully−extended and inter-mediate cable positions, and redundant switches. The hoist installation is mountedto the airframe by a support tube and strut assembly (Ref. Figure 10−17).

Hoist speed control is accomplished by a command applied to either the variablespeed switch on the hoist operator’s control pendant or the constant speed switchlocated on the collective control module. With the load selection switch set at 250,cable speed is 225 feet (68.5 meters) per minute. With the load selection switchset at 600, cable speed is 100 feet (30.5 meters) per minute (Ref. Figure 10−15).If the load select switch is set at 250 and the hoist load is above 250 LB, a flashingwarning light will illuminate and the hoist speed will automatically be reducedto 100 feet (30.5 meters) per minute.

The controller also passes cable position information from the hoist. This positioninformation is absolute and will continue to provide cable position information ifpower is interrupted.

The pilot’s payout switch overrides the hoist operator. When the pilot operates thepayout switch, the hoist is automatically set to the 600 LB 100 feet (30.5 meters)per minute mode.

Additional information pertaining to the hoist installation may be found in theBreeze−Eastern Corp. manual TD−92−015.





É

SUPPORTTUBE

STRUT

CONTROLPENDANT

VARIABLE SPEEDCONTROLLER

HOIST SUPPORT ASSEMBLY ROTATED

HOOK

F92−177A

HOIST ASSEMBLY(FAIRING

REMOVED)

FAIRING

MID SKID GUARD

AFT

INBOARD

WHITE STRIPE

15°15°

WHITE STRIPE

MID SKIDGUARD

FWD SKID TUBECOVER

AFT SKID TUBECOVER

HAND HOLD(OPTIONAL)

Figure 10−17. Rescue Hoist Installation






Table 10−6. Servicing Materials


Hoist assembly:

MIL−L−7808 Stauffer Jet I Stauffer Chemical Co.380 Madison AvenueNew York, NY 10017

American PQLubricant 6899

American Oil andSupply Co.

Mobil Avrex STurbo 256

Mobil Oil Co.

Brayco 880H Bray Oil Co1925 Marianna StreetLos Angeles, CA 90032

Exxon TurboOil 2389

Exxon Co.

Optional EquipmentRemovableCopilot Controls




10−11.OPERATING INSTRUCTIONS: REMOVABLE COPILOT CONTROLS

PART IGENERAL

The Removable copilot controls allows the aircrew to change their cockpit configura-tion from dual to single controls and back to dual, as desired, without the use oftools.

PART IILIMITATIONS

Flight crew:

Single pilot operation from the copilot seat is not approved with removable copilotcontrols installed.

Placards:

NO SINGLE PILOT OPERATIONUSING THIS CONTROL STICK

Figure 10−18. Collective and Cyclic Placards

PART IILIMITATIONS

No change.


Copilot cyclic stick removal (Ref. Figure 10−19):

Pull back hook tape fasteners (Velcro) and remove cyclic boot.

Detach P1 connector from receptacle on bulkhead.

Detach bonding jumper.

Remove quick release expandable bolts from from cyclic. Slide cyclic forwardto remove.

Remove protective cover or jumper plug from adjacent dummy receptacle andinstall it on J143.

Reinstall cyclic boot.





Properly stow cyclic.

Copilot cyclic stick installation:


NOTE: Verify correct operation of cyclic switches following installation.





COPILOT QUICK RELEASE

OPEN CLOSED

CYCLICBASE

CYCLIC STICK ASSEMBLY

BOOT

BONDING JUMPER

EXPANDABLE DIAMETER BOLT

BULKHEAD (REF)

COPILOT QUICK RELEASE

JUMPER PLUG OR PROTECTIVE COVER

P−1 CONNECTORJ143

DUMMY RECEPTACLE

F927−060

Figure 10−19. Removable Copilot Cyclic Control





Copilot collective stick removal (Ref. Figure 10−20):

Detach electrical connector P1 from receptacle J532 and connect it to dummyreceptacle after removing dust cap.

Pull back hook tape fasteners (Velcro) along collective boot.

Remove quick release pin by depressing button on top of pin and pull pin out.

Slide collective forward to remove.

Properly stow collective.

Copilot collective stick installation:


NOTE: Verify correct operation of collective switches following installation.

P1BOOT

QUICKRELEASE PIN

(

DUMMYRECEPTACLE

F927−061

J532

DUST CAP

Figure 10−20. Removable Copilot Collective Control






Use the weight information listed below to determine C.G. shift following controlremoval or installation.

ITEMWEIGHT

(LB)LONGITUDINAL

STATION(ARM)

LATERALSTATION

(ARM)

MOMENT(IN−LB)


Collective control 3.60 142.35 −27.60 512.46 −99.36

Cyclic 3.11 119.86 −15.70 372.77 −48.82

Cyclic boot 0.28 118.65 −15.85 33.22 −42.29

Pedal cover 0.56 101.14 −15.85 56.64 −8.88

Cyclic hole cover 0.45 120.00 −15.85 54.00 −7.13


MD900 (902 Configuration with PW 207E) Optional EquipmentAirframe Fuel FIlter

Original 10−55

FAA ApprovedReissue 1

10−12. OPERATING INSTRUCTIONS: AIRFRAME FUEL FILTER

PART IGENERAL

The airframe fuel filter incorporates a filter unit mounted in series between theaircraft fuel system and the engine fuel system. A pressure sensing switch in thefilter body will illuminate a FUEL FILTER L or R caution light on the instrumentpanel when the fuel differential pressure across the filter increases to a preset level.When the filter becomes fully clogged, a bypass valve contained in the filter unitopens and the fuel bypasses the filter element.

PART IILIMITATIONS

Placards:

When the airframe fuel filter is installed, the following placard is required:

AIRFRAMEFUEL FILTER

INSTL

USE PRIMARYFUELS ONLY

F927−094

Fuel system:

Fuel specification

Primary fuels only (Ref. Section II).

Optional EquipmentAirframe Fuel FIlter



Original10−56



FUEL FILTER CLOGGED

Indications: Yellow FILTERFUEL

L R

caution indicator ON.

Conditions: Left, right, or both airframe fuel filter(s) in impending bypass condition.

Procedures: With both fuel boost pumps operational.

� Continue the flight in progress.

� Service the airframe filter prior to next flight.

Procedures: With single fuel boost pump failure.


� Leave operating fuel boost pump ON.

If flight is continued into low fuel conditions (less than 300 poundsremaining), the engine with the failed boost pump will experiencean early flame out due to a loss of the fuel transfer system (Ref.Figure 10−21).

� Land as soon as practical if fuel level falls below 300 pounds.

Procedures: With dual fuel boost pump failure.

� Land as soon as possible.

FUEL BOOST PUMP FAILURE

Procedures:

� Leave operating fuel boost pump ON.

If flight is continued into low fuel conditions (less than 300 poundsremaining), the engine with the failed boost pump will experiencean early flame out due to a loss of the fuel transfer system (Ref.Figure 10−21).


� Land as soon as practical if fuel level falls below 300 pounds.

CAUTION

CAUTION


MD900 (902 Configuration with PW 207E) Optional EquipmentAirframe Fuel FIlter

Original 10−57



Preflight:

� Power switch BAT/EXT

� Left and right boost pumps ON

Attempting to drain the fuel filter with boost pumps off will causeair to enter the fuel system.

� Filter unit, drain valve, and associated lines CHECK FOR LEAKS

� Filter unit drain valve PRESS TO DRAIN

NOTE: If the aircraft has been exposed to freezing temperatures, failure of the drain maybe due to ice formation in the filter element.

� Press−to−test button PRESS

� AIRFRAME FILTER caution lights ON

� Left and right boost pumps OFF


Post Flight:

� When ambient temperature is expected to go below freezing, any water in the fil-ter unit should be drained following completion of flight.

� Following the completion of the flight in progress after illumination of the FUELFILTER caution light, service the filter prior to the next flight.

CAUTION

Optional EquipmentAirframe Fuel FIlter



Original10−58


LH DRAIN LINE

NOTE: LEFT HAND SHOWN; RIGHT HAND OPPOSITE. NOT TO SCALE.

F927−065

ELECTRICALCONNECTOR

FUEL FILTER ENCLOSURE

DRAIN HOSE

INLET HOSE

OUTLET HOSE

DRAIN VALVE

PRESS TO TESTBUTTON

VIEW LOOKING AFT

ENGINE DECK

RH FUEL FEED SYS

VENT OVBDDRAIN (2 PL) FUEL CELL

LH FUELFEED SYS

RIGHTENGINE

LEFTENGINE

AIRFRAMEFUEL FILTER

FUEL FILTER DRAIN (2PL)

FUEL BOOST PUMPS

FUEL TRANSFER SYSTEM

Figure 10−21. Airframe Fuel Filter Installation and Block Diagram

CSP−902RFM207E−1ROTORCRAFT FLIGHT MANUALMD900 (902 Configuration with PW 207E) Optional Equipment

SX−16 Night Sun: Aft Mount

Original 10−59


10−13. OPERATING INSTRUCTIONS: SX−16 NIGHT SUN WITH AFTMOUNT

PART IGENERAL

These operating instructions describe the SX−16 Nightsun Searchlight installationfor the MD900 as installed with the aft mount. Additional information may be foundin the SX−16 operations manual. Whenever the installation or operation of thisspecial mission equipment affects the operation of the basic helicopter, appropriatemention of the affected procedure, limitation, operation, will be described herein.

The searchlight is mounted to a trapeze−style, quick−disconnect, aft support mountattached to the rear of the fuselage (Ref. Figure 10−22)

PART IILIMITATIONS

A landing light shall be switched on when operating below 100ft AGL with thesearchlight (SX−16) on.

The use of the SX−16 as a landing light is not approved.

Do not turn ON the SX−16 while on the ground.


No change.


Ensure that the searchlight assembly has cooled before handling.

Do not turn the searchlight ON while on the ground.

Temporary blindness may occur to personnel if searchlight isaimed at vehicles or other aircraft at distances closer that 330 feet(100 meters).

In the infrared (IR) mode, the light beam is invisible and is a hazardto personnel at distances closer that 425 FT (130 meters). Do notperform operational checks of the searchlight with the IR filter inplace while helicopter is on the ground.

WARNING

Optional EquipmentSX−16 Night Sun: Aft Mount


Original10−60


Preflight checks:

� Aft mount (Ref. Figure 10−22):

� � Support tube and quick release pins CHECK ATTACHMENT� � Vertical and side struts CHECK ATTACHMENT

� � Bonding jumpers CHECK ATTACHMENT� � Mounting plate CHECK ATTACHMENT� Searchlight (Ref. Figure 10−23):

� � Drive assemblies CHECK� � IR filter CHECK CONDITION� � Clear lens CHECK CONDITION� � Cooling fan intake NO OBSTRUCTIONS� � Searchlight mounting CHECK ATTACHMENT

Search light operation (while airborne only):

� Turning searchlight ON:

� � Master switch ON; THEN TO START

Holding the start switch to START longer than 5 seconds will causedamage to the lamp.

NOTE: Do not operate searchlight while on ground unless conducting maintenancechecks. Ground checks may be accomplished with generator power or with theaircraft connected to a GPU.The magnetic compass may become inaccurate with the SX−16 ON.

• • Focus and directional control switches OPERATE ASNECESSARY

� Turning searchlight OFF:

Do not turn lamp OFF until lamp is fully illuminated.

� � Searchlight SET TO NEUTRALPOSITION

� � Master switch OFF

NOTE: Allow lamp to cool for one minute before turning back ON.

CAUTION

CAUTION



Original 10−61


F927−074A

BONDINGJUMPER

QUICKRELEASE PIN

AFT/SIDE MOUNT STRUT AND SUPPORT TUBE ATTACHMENTS

SX−16

VERTICALSTRUT

BONDINGJUMPER

QUICKRELEASE PIN

SIDE STRUT

SEARCHLIGHTSUPPORT TUBE

Figure 10−22. SX−16 Aft Mount InstallationBaggage compartment access:

NOTE: To gain access to the baggage compartment, the searchlight and supportassembly must be lowered. Lowering may be accomplished with generatorpower or with the aircraft connected to a GPU.

Before attempting to lower the searchlight and support, thesearchlight must be aimed toward the ground and the infrared lensmust be lowered. Damage to the lamp assembly and aircraft canoccur if lamp assembly is left in the up position while lowering.

CAUTION



Original10−62


Lowering searchlight and support (Ref. Figure 10−22):

� Hand controller

� � Master switch ON

� � IRCO switch CLOSE

� � Directional control switch DOWN

The searchlight must be aimed toward the ground and the infraredlens must be lowered. Damage to the lamp assembly and aircraftcan occur if lamp assembly is left in the up position while lowering.


� Eectrical master panel:

� � Power switch OFF

� SX 16 Aft mount:

� � Side strut bonding jumpers SEPARATE

� � Vertical strut bonding jumper SEPARATE

� � Side strut quick release pins REMOVE

� � Aft strut release pin REMOVE

NOTE: While holding the searchlight vertical support strut remove the quick release pinfrom the vertical strut, gently lower the SX−16 to the ground.

Raising searchlight and support (Ref. Figure 10−22):

Raising searchlight and support is opposite of lowering except there is no require-ment for electrical power.

CAUTION



Original 10−63


ÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌ

NOTE: IR FILTER SHOWN RETRACTED

IR FILTER DRIVEASSEMBLY

ELEVATION DRIVEASSEMBLY

AZIMUTH DRIVE ASSEMBLY

IR FILTER

CLEAR LENS

F927−063AHAND CONTROLLER

MASTERSWITCH

DIRECTIONALCONTROL

FOCUSCONTROL

IR LENSCONTROL


Master OFFON

START

Removes electrical power from SX−16 system.Turns on electrical power to SX−16 system.Energizes lamp starting circuit.

DirectionalControl

LEFT, RIGHT,UP, DOWN

Allows aiming of the searchlight.

FOCUS Press Momentary switch that changes light beam spread.

IRCO CLOSEOPEN

Positions IR filter in front of lamp.Retracts IR filter.

Figure 10−23. SX−16 Searchlight Assembly



Original10−64



No change.


ITEMWEIGHT

(LB)LONGITUDINAL

STATION (IN)(ARM)

LONGITUDINALMOMENT(IN−LB)

LATERALSTATION (IN)

(ARM)

LATERALMOMENT(IN−LB)

SX−16 Search light and sup-port assembly

66.0 279.3 18433.8 0.0 0.0


The SX−16 searchlight installation consists of a gimbal mounted searchlight assem-bly attached to an airframe searchlight mount, an electrical junction box a handcontroller. A 70 AMP circuit breaker, located on the baggage compartment circuitbreaker panel, recieves power from a generator bus.

The SX−16 has a 1600 Watt short arc Xenon lamp with a peak beam intensityof 30 million candlepower and a typical range of 3200 feet (1 km).


NOTE: A ground power unit should be used while conducting ground maintenancechecks of the SX−16.


RDR 1400C Weather Radar

Original 10−65


10−14. OPERATING INSTRUCTIONS: RDR−1400C WEATHER RADAR

PART IGENERAL

No change.

PART IILIMITATIONS

Ambient temperature limits:

Maximum temperature for operating radar is 40°C.

National safety regulations for radar operations shall be followed.

Do not operate radar within 50 FT (15M) of refueling vehicles orcontainers containing flammables or explosives.

Do not operate radar within 100 FT (30M) of refueling operations.

Do not allow personnel within 15 FT (5M) of area being scannedby antenna when system is transmitting.

Placard:

WEATHER RADAR SYSTEM MUST BE TURNED TO �STBY" OR �OFF" WHEN ON THE GROUND.


No change.


NOTE: Additional operating instructions may be found in the Telephonics Pilot’s Guide,RDR−1400C Color Weather and Search & Rescue Radar, publication number106501.

Preflight checks:

• Radome CHECK CONDITION

Engine starting:

Ensure that the RDR 1400C is not switched ON during engine starts.

Landing:

• Function switch SET TO “STBY”, OROFF

WARNING

Optional EquipmentRDR 1400C Weather Radar


Original10−66



No change.


RDR−1400C interface with EFIS 40 or EFIS 50:

The Telephonics Weather Radar RDR1400C is an integrated system, interfacingwith the LITEF LCR 92S AHRS and the Honeywell EFIS 40 installed by theHoneywell (formerly Allied Signal, Inc.) IFR STC SR00436WI−D, or theHoneywell EFIS 50 installed by the Heli−Dyne STC SR09151RC.

The Telephonics Weather Radar system consists of 5 flight−line replaceable units:

Receiver−Transmitter RT−1401BControl Panel CP−113Navigation Concentrator NC−104BRadar Antenna Drive unit DA−1203AAntenna Array AA−4510A

The Receiver−Transmitter (R/T) is installed in the nose of the MD900. TheR/T provides pulsed X−band output signal to the sector−scanning antenna. Thereflected signal is amplified by the receiver, digitized and then routed to theNavigation Concentrator for conversion to ARINC 453 data for the EFIS SymbolGenerators. Operating parameters permit optimum performance in each of thefive operational modes: 3 search, weather, and beacon modes

The Radar Control Panel is located in the center console. All of the WX systemcontrols except for display range are located on this control panel. Mode selections,Antenna Tilt, Search gain, BCN gain, and Scan are available for selection. TheAntenna Tilt and Roll adjustment screws are also located on this panel.

The Navigation Concentrator (NC) is installed in the nose, just above theR/T. It provides interface capabilities to convert WX R/T data into ARINC 453format required by the EFIS Symbol Generator. The Nav Concentrator receiveslow speed ARINC 429 data from the EFIS containing the range selected andconverts this into usable information for the R/T. The NC also receives low speedARINC 429 data from the Control Panel converting it into mode and controlcommands for the R/T. Lastly, the NC converts 429 attitude information fromthe AHRS for the R/T.

The Radar Antenna Assembly consists of a drive unit and a 10 inch array.The motor driven drive unit positions the array in azimuth and elevation axis.Scans are performed in 90 degree sectors. Stabilization is in accordance withpitch and roll signals from aircraft AHRS as converted by the NC−104B. Tiltis selectable for +/− 15 degrees from horizontal from the Control panel. The 10inch array has a range of 185NM.



Original 10−67


F927−078

ÉÉÉÉÉÉÉÉÉÉÉÉ

ATTITUDE

NAVCONCENTRATOR

RT−1401B

ANTENNA DRIVEDA−1203

ANTENNAAA−451A

CONTROL PANELCP−113

WAVE GUIDE

EFIS #1 EFIS #2

RANGE

ARINC 429

DATA

ARINC 453

ON/OFF

TILT

ROLL TRIM

ARINC 429

RANGE

ARINC 429

DATA

ARINC 453

ARINC 429

Figure 10−24. RDR 1400C EFIS System InterfaceSystem Function/Operational Modes:

Search modes

There are three search modes, with 6 ranges. Two of the search modes (SR1,SR2) can detect and display surface targets down to a minimum range of1000 feet. Even under adverse conditions, targets such as a small boat canbe detected. The search modes permit ground mapping or searching for topo-graphical features such as bodies of water, islands, high ground, bridges,etc. The slant range from the aircraft to a reflective ground object, withinthe area of scan, can be deduced directly from the display. The Search 1 modeuses special clutter rejection circuitry and is designed for short−range (ie0.5, 1,2,5,10 and 20NM) mapping of targets in a sea clutter environment.



Original10−68


Search 2 is designed for precision ground mapping with very high resolutionof short ranges (ie 20NM or less). Search 3 is designed for maximum clutterreturns as required for mapping of oil slicks. On long ranges (beyond 20NM),Search 1 and Search 2 are similar to Search 3.

Weather Avoidance Modes (Wx and WxA)

By means of a radar echo displayed on the EFIS, the system can furnishcontinuous enroute weather information relative to rain cloud formation, rain-fall rate, thunderstorms with moisture, and areas of icing conditions. Digitalcircuitry provides a means for determining the relative density of the rainfallareas. With the EFIS display the pilot can see storm areas in his flight pathand can also distinguish corridors of relative calm through the storms. Thesystem detects the strong returns from high density rainfall and convertsthem into red areas on the EFIS display. If the pilot changes the mode tothe weather alert (WxA) mode and the red area is beyond the range displayed,the TGT ALRT will flash.

Beacon Mode

In the beacon mode, the system can interrogate and receive pulses from afixed transponder(s) located within a range up to 160 NM. The coded repliesare received on a special beacon frequency (9310 MHz). The radar indicatordisplays beacon returns from both 2−pulse and DO−172 6 pulse transponders,located in range and bearing with respect to the aircraft.

The beacon mode can be operated alone, or combined with either the weatheror search modes.



Original 10−69


ONTST

STBY

OFF

60

SRCH GAIN

MIN MAX

BCN

GAIN

PUSH

CODE

MIN MAX

PULL

STABOFF

TILT0

+5

+10

+15

−5

−10

−15

Wx WxA

SRCH BCN

F927−090

FUNCTION SELECTOR

PRIMARY MODE SELECTORS

SECONDARY MODE SELECTOR ANDGAIN CONTROLS

ANTENNACONTROL

Figure 10−25. CP 113Operational controls:

Function Selector

OFF Removes system power.

STBY System is operationally ready; no display.

TEST Displays a test pattern without transmitting, identified by TEST andRT FAULT.

ON System transmits in normal operation.

60� is not supported by this installation and functions the same as �ON".

Primary Mode Selectors (PUSH ON/ PUSH OFF)

Wx Selects weather mode, the primary mode of operation. Weather displayedand Wx appear on screen. When pressed again, the weather mode is removed.If no other mode button is active, the Wx mode remains.

WxA Selects weather alert mode. WxA appears and Target Alert is enabled.

SRCH Pressing this push button selects the three Search modes in sequentialcyclic manner (i. e., Search 1, Search 2, Search 3, etc.). Search modes are as follows: Search 1 − Sea clutter rejection. Active on the ten− mile range or less. Search 2 − Short range precision mapping. Active on the ten− mile rangeor less. Search 3 − Normal surface mapping.

NOTE: Beacon mode is compatible with both weather mode and Search mode.



Original10−70


BCN Pressing this button will select the two Beacon type formats to be se-lected. Sequentially pressing the Beacon button will select the following bea-con modes: Beacon Only Mode − Beacon A, Beacon B, Beacon A, Beacon B, . . . Dual Mode (Beacon/ Weather or Beacon/ Search) Beacon A, Beacon B, BeaconOff, Beacon Am . . . Beacon Only Mode: − If the starting mode of operation is weather (or search),then pressing the beacon button will place the system in Beacon/ Weather(Search) mode. To activate beacon only mode, press the weather (or search)button to turn off weather (search) mode. To reactivate dual mode, press eitherthe weather or search buttons.

Beacon A − Standard 2−pulse beacon Beacon B − DO−172 compatible beacon

Secondary Mode Selector and Gain Controls

BCN GAIN − The Beacon Gain is a rotary potentiometer that controls thegain of the Beacon receiver.

SRCH GAIN − The Search Gain is a rotary potentiometer that controls thegain of the Search receiver.

CODE − Pressing this switch selects Beacon Codes in a sequential cyclic fash-ion (i. e., Code 0, Code 1, Code 2, . . . Code 15 or Code 0, Code 1, Code 2,. . . Code 9) depending on Beacon Mode selected. The selected code is annun-ciated on the MFD.

When DO−172 Beacon (Beacon Mode B) is selected via the BCN button, thetotal of sixteen codes (0−15) can be selected by the Code switch. Selectinga Standard two−pulse Beacon (Beacon Mode A) via the BCN button, the totalof ten codes (0−9) can be selected by the Code button. The Code button isnot active unless the Beacon mode has been selected.


LEO−II−A5 Observation System

Original 10−71


10−15. OPERATING INSTRUCTIONS: LEO−II−A5 OBSERVATION SYSTEM

PART IGENERAL

These operating instructions describe the LEO−II−A5 installed either on the �univer-sal mount" or the External Paload Mounting System (EPMS). Additional informationmay be found in the LEO−II−A5 operator manual.Whenever the installation or operation of this special mission equipment affectsthe operation of the basic helicopter, appropriate mention of the affected procedure,limitation, operation, will be described herein.

PART IILIMITATIONS

EPMS only:

Maximum airspeed 140 KIAS.

Do not use the equipment adaptors as steps.


No change.


NOTE: Additional operational procedures may be found in the Cumulus LEO−II−A5operating instructions manual.

Preflight checks:

• Stabilized turret assembly CHECK CONDITIONAND ATTACHEMENTTO MOUNTINGSURFACE

• EPMS

• • Main beam and cross tube attachment brackets CHECK CONDITIONAND INTEGRITY OFSTEP ASSEMBLY

• • Cross tube attachment brackets CHECK ATTACHMENTTO LANDING GEAR

• • Equipment adaptor CHECK CONDITIONAND SECURITY

Engine starting:

Ensure that the laptop control unit power switch is OFF during engine starts.

Optional EquipmentLEO−II−A5 Observation System


Original10−72


FRONT CROSS TUBE AT-TACH BRACKET

AFT CROSS TUBE AT-TACH BRACKET

EQUIPMENT ADAPTOR

EPMSMAIN BEAM

UNIVERSAL MOUNT

STABILIZED TURRETASSEMBLY

F927−064A

Figure 10−26. LEO−II−A5 Mounting


No change.


LEO−II−A5 Observation System

Original 10−73



ITEMWEIGHT

(LB)LONGITUDINAL

STATION (IN)(ARM)


LATERALSTATION (IN)

(ARM)


Universal Mount:

Stabilized turret assemblywith external harness

94.90 83.60 7933.60 0.0 0.0

EPMS:

Main beam 28.75 181.90 5229.60 +41 +1178.8

Leo−II Adaptor (fwd RH) 12.80 119.90 1534.70 +41 +524.80

Stabilized turret assemblywith external harness

94.90 105.50 10011.95 +41 +3890.9

Optional EquipmentLEO−II−A5 Observation System


Original10−74



The LEO−II−A5 Observation System consists of a Stabilised Turret Assembly (STA),Control Electronics Unit (CEU) and Laptop Control Unit (LCU).

Stabilized Turret Assembly:

The STA consists of three sub units namely the Stabilised Turret Unit (STU),Sensor Pack Assembly (SPA), and the Front Shell Assembly.

The STA may be mounted to either the universal mount or the EPMS (Ref.Figure 10−26).

Stabilised turret unit (STU)

The STU is a four−axis design, gyro stabilised in outer azimuth and elevationand inner azimuth and elevation. The STU provides the remote movementof the Sensor Pack Line−Of−Sight (LOS), corresponding to the centre of theimaged scene, to any position within the Field−Of−Regard (FOR). The FORfor the LEO−II−A5 Observation System is mechanical +20 degrees to –105degrees in elevation and a continuous 360 degrees in azimuth from thelook−ahead position.

Sensor pack Assembly (SPA)

The Sensor Pack contains the gyroscope, Sensor Control Electronics PCB,Triple QWIP thermal imager with three fields of view optics, TV Zoom CameraAssembly with a Three−CCD camera and 54X zoom lens, and a DummySpotter TV Assembly in the standard configuration.

Front Shell Assembly

The Front Shell Assembly consists of an aluminium shell into which windowsare fitted to accommodate the optical paths of the various sensors of the SensorPack Assembly. It also houses a desiccant holder, which eliminates moisturebuilt−up in the STA.

CONTROL ELECTRONIC UNIT (CEU):

The CEU houses all computing and control functions of the STA. It is the centralpoint of transfer for system signals, command, and data I/O. The CEU controlspower distribution and system timing. It implements software algorithms thatperform such tasks as the generation of overlay symbology and controlling themenu−driven functions of the system. The CEU communicates with other systemmodules through a serial data bus.

LAPTOP CONTROL UNIT (LCU)

The LCU is a lightweight laptop held unit by which control of the ObservationSystem is executed.

The Laptop Control Unit is held in place with a neck strap on the operator’slap during flight. The Laptop Control Unit is used by the operator to steer theSTA and control the sensor fitted in the Sensor Payload Assembly.


Annunciator Panel

Original 10−75/(10−76 blank)


10−16. OPERATING INSTRUCTIONS: ANNUNCIATOR PANEL

PART IGENERAL

This panel advises the pilot of the operating mode and status of auxiliary fuel trans-fer, pitot heat, ground power unit door (aft mounted battery only), and airframefuel filter. These annunciator lights may be installed individually or in part as deter-mined by the optional equipment installed and are only available with the HoneywellIFR Avionics STC.


For actions following illumination of any of the caution/advisory annunciators, referto Section X for Supplemental Fuel System and Airframe Fuel Filter; for the remain-ing caution/advisories, refer to Section III.


To check the operation of the annunciators, press the ANNUN test button on theAvionics switch panel.

F927−110

AHRS1 AHRS2 TEST

DG

CCW

CW

MRKR AP/SAS ANNUN

SLVD

DG

SLVD

CCW

CW

DMEHOLD

ANNUNDIM

BRIGHTRELEASE

Figure 10−27. Caution and Advisory Annunciators


Moving Map Navigation Systems

Original 10−77


10−17. OPERATING INSTRUCTIONS: MOVING MAP NAVIGATIONSYSTEMS

PART IGENERAL

Whenever the installation or operation of this special mission equipment affectsthe operation of the basic helicopter, appropriate mention of the affected procedure,limitation, or operation, will be described herein.

EuroNav III:

The EuroNav III is a GPS based navigation system with moving map displayand integrated task management system. The system is protected by a 5 amperecircuit breaker located on the baggage compartment circuit breaker panel andis approved with for use with the AI−500 monitor.

AeroNav� II:

The AeroNav II is a Moving Map and Task Management system. The systemis protected by a 5 ampere circuit breaker located on the baggage compartmentcircuit breaker panel and is approved with for use with the AVM4090 monitor.

Dornier DKG4:

The DKG 4 moving map is a digital map system that provides pilot navigationsupport. The digital pictorial information is displayed on the Avalex AVM4090Monitor in the cockpit and/or on a Skyquest monitor in a cabin workstation(if installed). The system is protected by a 2 ampere circuit breaker located onthe baggage compartment circuit breaker panel.

PART IILIMITATIONS

The use of any approved moving map navigation system as a primary means ofnavigation is not approved.

EuroNav III:

Normal ambient temperature range is 0°C to 40°C for cockpit mounted AI−500monitor.

Optional EquipmentMoving Map Navigation Systemss


Original10−78



AeroNav� II and Dornier DKG4:

Refer to AeroNavII/III HB−NAV−350, and Dornier DKG4 HB103−000000B/01operator manuals for operational information.

F927−113

PWR CMPT VIDEO FRZ/Z

Figure 10−28. AVM4090 Display Controls

Table 10−7. AVM4090 Display Control FunctionsControl Function

PWR Turns unit on or off when pressed.

CMPT Displays the video source. Each time this button is pressed, the video source(VGA1 or VGA2) appears on the lower left side of the display.

VIDEOThis button allows the display of FLIR, CCD, or VCR playback. Each time this but-ton is pressed, the video source (VID1, VID2 or VID3) appears on the lower leftside of the display.

FRZ/Z This button processes video and CMPT (SVGA) inputs and either freezes or zoomsor toggles between the two.

These buttons are used to control display brightness and operates the menu sys-tem. The monitor retains the previous brightness setting upon restart.

NOTE: Refer to the Avalex AVM4090 Operations guide for more detailed information.

EuroNav III:

Refer to Table 10−8 for AI−500 monitor operating instructions.


Moving Map Navigation Systems



F927−109

Figure 10−29. AI−500 Monitor

Table 10−8. AI−500 Control FunctionsControl Function

POWER When the Power button is depressed to ON the adjacent power indicator lights togreen. To turn the power off, depress the POWER button again.

VIDEO SELECT The two buttons, identified as IR/CCD or MAP select their respective video source.The selected button lights in green when depressed.

MENU Pressing the MENU button provides a screen menu which allows the operator toadjust the following levels: CONTRAST: (Picture Contrast) Adjusts the picture contrast. Decreasing the valuelowers the contrast, and increasing the value raises it. Adjustable range: −30 to+30

BRIGHTNESS: (Picture Brightness) Adjusts the picture brightness. Decreasingthe value makes the picture darker, and increasing the value makes it brighter.Adjustable range: −30 to +30

SHARPNESS: (Picture Sharpness) Adjust the picture sharpness. Decreasing thevalue makes the picture softer, and increasing the value makes it more sharp. Ad-justable range: −30 to +30

CHROMA: (Picture Chroma) Adjusts the picture chroma. Decreasing the valuemakes the picture lighter, and increasing the value makes it deeper. Adjustablerange: −30 to +30

PHASE: (Picture phase, NTSC only) Adjusts the picture phase. Decreasing the val-ue makes the picture more reddish and increasing the value makes it more green-ish. Adjustable range: −30 to +30

COLOUR SYSTEM: (Colour System) Displays the colour system (NTSC or PAL)used by the video equipment.

POWER SAVE: (Power Save) With the power save function set to ON (active) themonitor automatically enters standby mode when no video signal is input. Whenthe power save function is active the POWER indicator blinks in green. When avideo signal is input, the power save function becomes inactive and the monitor isrestored to normal operation. Pressing any buttons on the front operation panelalso sets the power safe function to inactive.

The power save function becomes active when no video signal is input for over 30seconds.

COLOUR SWITCH: Turns the picture into black and white for checking the whitebalance.

The MENU + and − buttons control the settings for each of these controls.


W.E.S.T. Battery Protection System

Original 10−81


10−18. OPERATING INSTRUCTIONS: W.E.S.T BATTERY PROTECTION SYSTEM

PART IGENERAL

The W.E.S.T. Ensave 02 battery protection system allows the pilot to use selectedelectrical equipment with the aircraft battery OFF.


No change.


Operating the Ensave 02:

NOTE: The battery protection system will not function with generator/GPU power andis designed to be operated with the POWER switch OFF.

Pressing the indicator button, activates the battery protection system.

Press the indicator button once to activate the GND 1 consumers.

Press the indicator button a second time to activate and add the GND 2 con-sumers.

Pressing the indicator button a third time deactivates the battery protectionsystem.

Once the system is activated, battery power to GND 1/GND2 consumers willbe available for one hour. After the one hour period is completed, a buzzer soundsadvising the operator that battery power to the GND 1/GND 2 consumers willbe removed within 10 seconds.

NOTE: The operator may reactivate the battery protection system following anautomatic shut off by pressing the button again, however sufficient battery powerfor starting may not be available.

F927−097

GND 1

GND 2

Figure 10−30. Ensave02 Indicator/Button

Optional EquipmentW.E.S.T. Battery Protection System


Original10−82



This system uses a timer and �consumer" button functions to time the usage ofselected equipment for specific periods without draining the battery to the extentthat a battery start could not be accomplished. At the end of the specified time,a buzzer sounds to alert the pilot that battery usage has reached a defined point.


SX−16 Night Sun: EPMS Mount

Original 10−83


10−19. OPERATING INSTRUCTIONS: SX−16 NIGHT SUN WITH EPMSMOUNT AND LASER POINTER

PART IGENERAL

These operating instructions describe the SX−16 Nightsun Searchlight with laserpointer installed on the external payload mounting system (EPMS). Additional infor-mation may be found in the SX−16 operator manual and the W.E.S.T. �Operatorpanel for SX−16 and laser system". Whenever the installation or operation of thisspecial mission equipment affects the operation of the basic helicopter, appropriatemention of the affected procedure, limitation, operation, will be described herein.

PART IILIMITATIONS

A landing light shall be switched on when operating below 100FT AGL with thesearchlight (SX−16) on.

The use of the SX−16 as a landing light is not approved.

Do not turn ON the SX−16 while on the ground.

Maximum airspeed 140 KIAS.Do not use the equipment adaptors as steps.Searchlight to be in maximum up position and facing forward when not in use.


No change.


Ensure that the searchlight assembly has cooled before handling.

Do not turn the searchlight ON while on the ground.

Temporary blindness may occur to personnel if searchlight isaimed at vehicles or other aircraft at distances closer that 330 feet(100 meters).

In the infrared (IR) mode, the light beam is invisible and is a hazardto personnel at distances closer that 425 FT (130 meters). Do notperform operational checks of the searchlight with the IR filter inplace while helicopter is on the ground.

Avoid aiming the laser pointer at personnel below a distance of 72feet (22 meters).

WARNING

Optional EquipmentSX−16 Night Sun: EPMS Mount


Original10−84


PREFLIGHT CHECKS

� EPMS (Ref. Figure 10−31):

� � Main beam and cross tube attachmentbrackets

CHECK CONDITION ANDINTEGRITY OF STEP ASSEMBLY

� � Cross tube attachment brackets CHECK ATTACHMENT TOLANDING GEAR

� � Equipment adaptors CHECK CONDITION ANDSECURITY

F927−088A

ÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌ

NOTE: IR FILTER SHOWN RETRACTED

IR FILTER DRIVEASSEMBLY

ELEVATION DRIVEASSEMBLY

AZIMUTH DRIVE ASSEMBLY

IR FILTER

CLEAR LENS

HEAT SHIELD

MAIN BEAM

FRONT CROSS TUBEATTACH BRACKET

AFT CROSS TUBEATTACH BRACKET

AFT EQUIPMENT ADAPTOR

FWD EQUIPMENTADAPTOR

LASER POINTER

SX−16−FWD MOUNT

SX−16−AFT MOUNT

Figure 10−31. SX−16 EPMS Installation



Original 10−85


� Searchlight (Ref. Figure 10−31):

� � Drive assemblies CHECK

� � IR filter CHECK CONDITION

� � Clear lens CHECK CONDITION

� � Cooling fan intake NO OBSTRUCTIONS

� � Laser pointer CHECK ATTACHMENT

� � Searchlight mounting CHECK ATTACHMENT

� � Heat shield CHECK ATTACHMENT

AFTER TAKEOFF

The SX−16 is fully controllable in azimuth up to 85 KIAS. Above 85 KIAS, controlof the SX−16 may be affected by the slipstream, resulting in the searchlight movingon its own by slipping the clutch.

With the IR lens in the forward and horizontal position, the SX−16 can be controlledup to 80 KIAS . As the light is moved to the side and rear position, the lens isless affected by the slipstream and can be raised and lowered at progressively higherairspeeds up to 100 KIAS.

Search light operation (while airborne only):

� Turning searchlight ON:

� � Master switch ON; THEN MOMENTARILY TOSTART POSITION

NOTE: Do not operate searchlight while on ground unless conducting maintenancechecks. Ground checks may be accomplished with generator power or with theaircraft connected to a GPU.The magnetic compass may become inaccurate with the SX−16 ON.

• • W.E.S.T. control panel switches OPERATE ASNECESSARY

� Turning searchlight OFF:

Do not turn lamp OFF until lamp is fully illuminated.

� � Searchlight MAXIMUM UPPOSITION


NOTE: Allow lamp to cool for one minute before turning back ON.

CAUTION

Optional EquipmentSX−16 Night Sun: EPMS Mount


Original10−86


F927−114

INFRARED FILTERINDICATOR

SLAVEINDICATOR

LASERINDICATOR

SX−16 �ON"INDICATOR

GIMBAL DIRECTIONCONTROL


SX16 OffOf

Start

Removes electrical power from SX−16 system.Turns on electrical power to SX−16 system.Energizes lamp starting circuit.

FOCUS Press Momentary switch that changes light beam spread.

GimbalDirectional

Control

LEFT, RIGHT,UP, DOWN

Allows aiming of the searchlight.

IFCO UP Positions IR filter in front of lamp.Retracts IR filter.

Slave UP Connects to AeroNav slaving unit (or other system asdetermined by installation requirements).

Laser OffOn

Flash

Turns laser off.Turns laser on.Flashes the laser beam.

BLT UP Allows testing of laser system before flight or duringmaintenance.

Sync. UP Provides synchronization between the FLIR and SX−16zero−position.

Figure 10−32. W.E.S.T. SX−16 Control Panel






No change.


ITEMWEIGHT

(LB)LONGITUDINAL

STATION (IN)(ARM)


LATERALSTATION (IN)

(ARM)


Main beam 28.75 181.90 5229.60 −41 −1178.80

SX−16 Adaptor (aft LH) 8.75 251.10 2197.10 −41 −358.80

SX−16 Search light (aft LH) 47.50 264.80 12578.00 −41 −1947.50

SX−16 Adaptor (fwd LH) 8.75 122.00 1067.50 −41 −358.80

SX−16 Search light (fwd LH) 47.50 110.00 5225.00 −41 −1947.50


The SX−16 searchlight installation consists of a gimbal mounted searchlight assem-bly attached to an airframe searchlight mount, an electrical junction box a panel−mounted controller. A 70 AMP circuit breaker, located on the baggage compartmentcircuit breaker panel, receives power from a generator bus.

The SX−16 Control Panel incorporates full authority over the Nightsun system in-cluding an enhancement to automate beam �cool down" following use. The Panelalso incorporates full authority over the laser spotter device. Additionally, the laserpointer is coupled to the radar altimeter and automatically turns off when a predeter-mined decision height has been reached. If desired, the pilot/copilot may allow theSX−16 to operate in a synchronized mode with the FLIR LEO II surveillance systemif installed.

The SX−16 has a 1600 Watt short arc Xenon lamp with a peak beam intensityof 30 million candlepower and a typical range of 3200 feet (1 km).


NOTE: A ground power unit should be used while conducting ground maintenancechecks of the SX−16.


Smoke Detector

Original 10−89


10−20.OPERATING INSTRUCTIONS: SMOKE DETECTOR

PART IGENERAL

No change.

PART IILIMITATIONS

No change.


CARGO COMPARTMENT FIRE/SMOKE

Indications: Smoke detector warning tone in headset.


Procedures:






Procedures:


� AC/VENT switch VENT LOW OR VENT HIGH

� Cockpit door vents OPEN


Optional EquipmentSmoke Detector


Original10−90


� After landing:



PRE FLIGHT CHECKS: ELECTRICAL POWER ON

Baggage compartment:

� Circuit breaker panel cover REMOVE� SMOKE DET press−to−test button PRESS� Listen for smoke detector warning tone in

headset.CHECK

NOTE: A second crew member is required to perform this check.

� Circuit breaker panel cover REPLACE

RH REAR FUSELAGESHELL ASSEMBLY

SMOKEDETECTOR

F927−118

RIGHT AVIONICS BUS


ADF1 FMCTRL

FM1RT

FM2RT

FM3RT

DME STORMSCOPE

CAMERA NAV 1 RMI

PRESS TOTEST

SMOKEDET

BAGGAGE COMPARTMENT-CIRCUIT BREAKER PANEL

Figure 10−33. Smoke Detector and Press−To−Test Switch Location.


Smoke Detector




No change.


The smoke detector is a photoelectric device specifically developed for aircraft cargobay applications and is located on the upper right hand wall of the baggage compart-ment adjacent to the baggage compartment door. The detector incorporates specificdesign features that virtually eliminate the reliability problems typically associatedwith aircraft smoke detectors. The detector is a dual−channel, ratio−comparing de-vice in which one channel detects the presence of smoke and the second channelserves as a reference. By comparing smoke and reference ratios, the detector isable to operate reliably despite dust, moisture, temperature changes, and aging.

The detector provides an alarm signal (sweeping tone) to the aircraft ICS systemwhen the output from the smoke channel exceeds a predetermined ratio to the outputfrom the reference channel. The warning tone is heard through the headset. A testinput activates a complete through−the−lens check of electronic and optical func-tions. The press−to−test button is located on the lower right hand corner of theright avionics bus.

The smoke detector system receives power from the battery bus and is protectedby a 5 amp. circuit breaker.


Crew Door Modification

Original 10−93


10−21.OPERATING INSTRUCTIONS: CREW DOOR MODIFICATION WITHQUICK RELEASE MECHANISM

PART IGENERAL

The quick release mechanism allows removal of crew doors in the field.

PART IILIMITATIONS

No change.


No change.


PRE FLIGHT CHECKS

� Left/Right crew door interior (Ref. Figure 10−34):

� � Door release handle up CHECK

� � Pull rods engaged in fork assemblies CHECK

Optional EquipmentCrew Door Modification


Original10−94


Cockpit door removal:

Open cockpit door.

Pull lower end of gas strut (1) up away from pivot post ball fitting (2) to remove.

NOTE: Fit between lower socket end of strut and ball fitting is by interference. Removalof strut from its attachment requires a snap action motion to pull away the socketend from the post ball fitting.

Push door release handle (3) forward to disengage pull rod assemblies (10) andrelease cockpit door retaining bosses (4).

NOTE: When activating cockpit door release mechanism, spacer, washer, and retainingbosses (6, 5, 4) will fall loose as the pull rods (10) are retracted. Place looseattaching hardware in a suitable location for reinstallation

Remove complete door assembly with retainers (7) in place on fork (8) assemblies.

Cockpit door installation:

NOTE: Refer to Figure 10−34 for proper placement of attaching hardware.

Install cockpit door assembly into door frame with retainers and spacer (7, 5)in place on fork assemblies (8).

Install washer (6), door restraint (9), and retaining boss (4) to lower fork assembly(8).

Install spacer (5) and retaining boss (4) on upper fork assembly (8).

Align pull rods (10) with retaining bosses (4).

Pull door release handle (3) backward and verify insertion of pull rods into retain-ing bosses.

Push down lower end of gas strut (1) onto lower pivot post ball fitting (2) toattach.

Close cockpit door.


Crew Door Modification



F927−130

DOOR RELEASE HANDLE

DOOR RESTRAINING STRAP

LOWER PULL ROD ASSEMBLY

UPPER PULL ROD ASSEMBLY

DOOR ATTACHING HARDWARE

DOOR RELEASE MECHANISM

1

2

3

45

6

78

9

10

GAS STRUT

2

4

10

5

87

Figure 10−34. Cockpit Door Attachment

Category A Operations


FAA ApprovedReissue 1Original 11− i

S E C T I O N XICATEGORY AOPERATIONS

TABLE OF CONTENTS

PARAGRAPH PAGEPart I General 11−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−1.1. General 11−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−1.2. Definitions − Category A Takeoff 11−1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−1.3. Definitions − Category A Landing 11−3. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Part II Limitations 11−5. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11−2.1. Clear Airfield, Heliport and Elevated Helipad 11−5. . . . . . . . . . . . . . . . . . . . . .

Figure 11−2.1. Takeoff and Landing Wind Azimuth Limitations 11−6. . . . . . . . . . . .

11−2.2. Maximum Takeoff and Landing Weight Limits 11−6. . . . . . . . . . . . . . . . . . . . .

Figure 11−2.2. Weight Altitude Temperature Limits − Clear Airfield 11−7. . . . . . . .

Figure 11−2.3. Weight Altitude Temperature Limits − Heliport/Elevated Helipad 11−8. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Part III Takeoff and Landing Procedures 11−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11−3.1. Clear Airfield Takeoff Procedures 11−9. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−3.1. Takeoff Timing Indicator Lights and Switch 11−9. . . . . . . . . . . . . . . .

Figure 11−3.2. Normal Takeoff and Takeoff Path 11−10. . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−3.3. Category A Rejected Takeoff 11−11. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−3.4. Continued Takeoff 11−12. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−3.2. Heliport/Elevated Helipad Takeoff Procedures 11−13. . . . . . . . . . . . . . . . . . . . .

Figure 11−3.5. Normal Takeoff Profile − Heliport/Elevated Helipad 11−13. . . . . . . . .

11−3.3. Landing Procedures − Clear Airfield, Heliport and Elevated Helipad 11−15.

Figure 11−3.6. Normal Landing Profile 11−15. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−3.7. Balked Landing 11−16. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−3.8. Continued Landing 11−17. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−3.4. Equipment Malfunctions 11−18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

IIDS Failure 11−18. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Category A Operations


11− ii

PARAGRAPH PAGEPart V Performance Data 11−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.1. Takeoff Performance 11−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.2. Takeoff Distance Required 11−19. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−5.1. Distance Required to Clear a 35 FT (11M)

Obstacle on Takeoff (Clear Airfield) 11−20. . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−5.2. Rejected Takeoff Distance Required (Clear Airfield) 11−21. . . . . . . . .

Figure 11−5.3. Distance Required to Clear a 35 FT (11M)

Obstacle on Takeoff Heliport/Elevated Helipad 11−22. . . . . . . . . . . . . .

11−5.3. Continued Takeoff FLight Path 11−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−5.4. OEI Takeoff Flight Path 11−23. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−5.5. Takeoff Distance Segment I − Distance Required to Climb from 35 FT (11M) to 200 FT (61M) HAT 11−24. . . . . . . . . . . . . . .

Figure 11−5.6. Acceleration Distance − Distance to Accelerate from

45 KIAS (83 KM/H) to VY at 200 FT (61M) HAT 11−25. . . . . . . . . . . . .

Figure 11−5.7. Takeoff Distance Segment II − Distance Required to

Climb from 200 FT (61M) HAT to 1000 FT (305M) 11−26. . . . . . . . . . .

11−5.4. Landing Performance − Open Airfield 11−27. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−5.5. Landing Performance − Heliport/Elevated Helipad 11−27. . . . . . . . . . . . . . . . .

Part IX Additional Operations 11−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11−9.1. Category A OEI Training 11−29. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 11−9.1. Recommended OEI Training Weight 11−30. . . . . . . . . . . . . . . . . . . . . . .

Category A OperationsGeneral




SECTION XICATEGORY A OPERATIONS

PART IGENERAL

11−1.1. GENERAL

Information contained in this section pertains to Category A operations only andsupplements information that appears in Sections 1 thru 10 of this manual.There are several ‘‘Parts’’ to this section. Each ‘‘Part’’ is associated with a respectiveSection from the RFM, with the exception of Part III which covers both normaland emergency procedures.

NOTE: Performance charts in this section are based on an aircraft with IPS, NACA inletdoor closed, and a power assurance check with zero margin for both engines.

11−1.2. DEFINITIONS − CATEGORY A TAKEOFF

Category A takeoff:

The takeoff must be performed in such a manner that in the event of a singleengine failure the helicopter must be able to:

Prior to TDP, return to, and stop safely on the takeoff area (rejected takeoff).

After TDP, continue the takeoff and climbout, and attain a configuration andairspeed that allows continued flight.

Take−off Decision Point (TDP):

Clear airfield

The TDP is a point that occurs 8 seconds after the takeoff procedure is initiated.The takeoff light will display a yellow ‘‘NO−GO’’ indication for 8 seconds.

The green ‘‘GO’’ indicator illuminates after the TDP.

Heliport/Elevated helipad

The TDP is a point 100 FT (30m) HAT and approximately 300 FT (91m)behind the center of the heliport.

HAT

Height above touchdown.





Rejected takeoff distance:

The horizontal distance required to land and come to a complete stop shouldthe engine fail before reaching TDP.

Continued takeoff distance:

The continued takeoff distance is the horizontal distance along the takeoff pathfrom the start of the takeoff to the point at which the rotorcraft attains andremains at least 35 feet (11m) above the takeoff surface, attains and maintainsa speed of at least VTOSS, and establishes at least a 100 ft/min (30m/mini) rateof climb, assuming the recognition of a critical engine failure at TDP.

Takeoff Segment distances:

Segment I distance:

The horizontal distance required to climb at VTOSS from 35 FT (11m) AGLto 200 FT (61m) AGL.

Acceleration distance:

The horizontal distance required to accelerate from VTOSS at 200 FT (61m)to VY.

Segment II distance:

The horizontal distance required to climb at VY from 200 FT (61m) AGLto 1000 FT (305m) AGL.

Take−Off Safety Speed (VTOSS)

The speed (40 KIAS [74km/h]) at which a safe take−off can be continued followingan engine failure.

Best rate of climb speed (VY):

The best rate of climb speed is that airspeed that achieves the best rate of climbat a given density altitude (Ref. Section V).





11−1.3. DEFINITIONS − CATEGORY A LANDING

Category A landing:

The landing must be performed in such a manner so that if the critical enginefails at any point in the approach path, the helicopter must be able to:

Prior to LDP, climb out and attain an airspeed that allows continued flight(balked landing).

After LDP, land and stop safely.

Landing Decision Point (LDP):

The landing decision point is the last point in the approach and landing pathat which a balked landing can be accomplished with the critical engine failedor failing and with the engine failure recognized by the pilot. This point is definedas 100 FT (30m) HAT and 35 KIAS (65km/h).

Landing Distance:

Clear Airfield

The horizontal distance required to land and come to a complete stop froma point 50 feet (15m) above the landing surface.

Heliport/Elevated helipad

The horizontal distance required to land and come to a complete stop froma point 25 feet (8m) above the landing surface.

Category A OperationsLimitations




PART IILIMITATIONS

NOTE: Limitations contained in this part pertain to Category A operations only andsupplements information that appears in Section II.

11−2.1. CLEAR AIRFIELD, HELIPORT AND ELEVATED HELIPAD

Flight Restrictions:

Aircraft equipped with Bendix/King KFC900 Flight Control System:

NOTE: The following information supersedes applicable limitations found in Bendix/KingIFR Avionics/KFC 900 RFMS 006−00845−0000 and 006−00845−0004 for STCSR00436WI−D.

For VFR flights at gross weights between 6251 (2835kg) and 6500LB(2948):

Maximum airspeed with autopilot engaged is 100 KIAS (185 km/h)Maximum Operating Altitude with autopilot engaged 5000 FT (1524m) HD

IFR flights are only approved at gross weights up to 6250 LB and:Autopilot must be operational.Maximum airspeed with autopilot engaged is 100 KIAS (185 km/h)Maximum Operating Altitude with autopilot engaged 5000 FT (1524m) HD

Environmental operating conditions:

Kinds of operations

This rotorcraft is certified in the normal helicopter category for day and nightVFR Category A operations when the appropriate instruments and equipmentrequired by the airworthiness and/or operating rules are approved, installedand are in operable condition.

Critical wind azimuth

Refer to Figure 11−2.1.

Weight altitude temperature limits

Open field: Maximum weight for Category A operations is 6500 LB (2948kg)or less as determined by Figure 11−2.2.

Heliport/Elevated helipad: Maximum weight for Category A operations is6500 LB (2948kg) or less as determined by Figure 11−2.3.

Maximum altitude for Category A operations is 10,000 FT (3048m) HD.

Power assurance checks:

Each engine must pass a power assurance check prior to takeoff (Ref. SectionV).

Heliport/Elevated helipad requirements:

Heliport/Elevated helipad restricted to a solid surface.

Minimum Heliport/Elevated helipad dimensions: 50 FT X 50 FT (15m X 15m).





ÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌÌ

F927−040

CATEGORY A TAKEOFFSAND LANDINGS WITH

WINDS FROM THE CROSSHATCHED AREA ARE NOT

PERMITTED.

30°330° 0°

Figure 11−2.1. Takeoff and Landing Wind Azimuth Limitations

11−2.2. MAXIMUM TAKEOFF AND LANDING WEIGHT LIMITS

Description: These charts show the maximum gross weight for a given temperatureand altitude for Category A operations from a clear airfield (Ref. Figure 11−2.2)or Heliport/Elevated helipad (Ref. Figure 11−2.3).

Use of Chart: The following example explains the correct use of the chart inFigure 11−2.2.

Example:Wanted: Maximum gross weight for Category A operations from a clear

airfield.Known: Outside air temperature = 28°CKnown: Pressure altitude = 2000 FT (610m)Method: Enter bottom of chart at 28°C. Move up to the 2000 FT (610m)

line and then directly to the left to read 6160 LB (2794kg).





F927−034−1D−20 −15 −10 −5 0 5 10 15 20 25 30 35 40 45 50 55

LIMIT

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

OAT − °C

GR

OS

S W

EIG

HT

− L

BS

(kg

)


10000 HD

LIMIT

NOTE: CABIN HEAT AND AC OFF

11000(3353M)

10000(3048M)

8000(2438M)

7000(2134M)

6000(1829M)

5000(1524M)

4000(1219M)

3000(914M)

2000(610M)

1000(305M)

SL −1000(−305M) PRESSURE

ALTITUDE − FEET

9000(2743M)

MAXIMUM OAT

(3048M)

(2268kg)

(2313kg)

(2359kg)

(2404kg)

(2449kg)

(2495kg)

(2540kg)

(2585kg)

(2631kg)

(2676kg)

(2722kg)

(2767kg)

(2812kg)

(2858kg)

(2903kg)

(2948kg)

(METERS)

Figure 11−2.2. Weight Altitude Temperature Limits − Clear Airfield





F927−034−2D

−20 −15 −10 −5 0 5 10 15 20 25 30 35 40 45 50 55

OAT − °C


NOTE: CABIN HEAT AND AC OFF

SL

MAXIMUM OAT

10000 HD

LIMIT

(3048M)

5000

5100

5200

5300

5400

5500

5600

5700

5800

5900

6000

6100

6200

6300

6400

6500

(2268kg)

(2313kg)

(2359kg)

(2404kg)

(2449kg)

(2495kg)

(2540kg)

(2585kg)

(2631kg)

(2676kg)

(2722kg)

(2767kg)

(2812kg)

(2858kg)

(2903kg)

(2948kg)

GR

OS

S W

EIG

HT

− L

BS

(kg

)

4000(1219M)

3000(914M)

2000(610M)

1000(305M)

−1000(−305M)


(METERS)

10000(3048M)

8000(2438M)

7000(2134M)

6000(1829M)

5000(1524M)

9000(2743M)

Figure 11−2.3. Weight Altitude Temperature Limits − Heliport/Elevated Helipad

Category A OperationsTakeoff and Landing Procedures




PART IIITAKEOFF AND LANDING PROCEDURES

NOTE: This section contains both normal and emergency procedures for Category Atakeoffs and landings as well as special procedures for equipment malfunctions.

11−3.1. CLEAR AIRFIELD TAKEOFF PROCEDURES

Takeoff timer operation and check:

To start the takeoff timer, push up on the TAKEOFF TIMER switch (Ref.Figure 11−3.1). This action turns on the yellow NO−GO light. After 8 seconds,the green GO light illuminates.

To shutoff the takeoff timer, push down on the TAKEOFF TIMER switch a secondtime.

The TAKEOFF TIMER switch is not functional on the copilot’s collective (if dualcontrols are installed).

NOTE: The GO and NO−GO lights dim when the LIGHT MASTER switch is placed inthe ON (‘‘night mode’’) position.

Follow the above procedure prior to performing a clear airfield takeoff.

NO−GO GO

YELLOWNO−GO LIGHT

GREEN GO LIGHT

F927−035

TAKEOFF TIMERSWITCH

Figure 11−3.1. Takeoff Timing Indicator Lights and Switch

NOTE: TDP is a point that occurs 8 seconds after the takeoff procedure is initiated. The takeoff light will display a yellow ‘‘NO−GO’’ indication for 8 seconds. At the TDP, the green ‘‘GO’’ indicator illuminates.





Normal takeoff and takeoff path:

� Power assurance check PASS� Takeoff timer CHECK� Pre takeoff check PERFORM� Hover ESTABLISH − 3.5 FT (1M) SKID HEIGHT; NOTE

HOVER TORQUE� Takeoff/climb SIMULTANEOUSLY:

START LEVEL ACCELERATION TAKEOFF(APPROXIMATELY 12° NOSE DOWN).START TAKEOFF TIMER, SET COLLECTIVE TO A TORQUE 10% ABOVEHOVER POWER.

AS AIRCRAFT PASSES THROUGH ETL MAINTAINPITCH ATTITUDE TO ALLOW CLIMB ANDCONTINUED ACCELERATION TO TDP. (AT TDP THE ALTITUDE SHOULD BEAPPROXIMATELY 20 FT (6M)HAT.)

NOTE: Category A timer will display NO GO for 8 seconds; then GO.

� After TDP CLIMB AND ACCELERATE TO VY.� Takeoff timer TURN OFF AS DESIRED

8 SECONDS

CLIMB AT VY

ACCELERATETO VY

TDP

HOVER AT 3.5 FT (1M)SKID HEIGHT

F927−041

Figure 11−3.2. Normal Takeoff and Takeoff Path





ENGINE FAILURE BEFORE TDP

Indications: Normal engine failure indications (Ref. Section III).

Conditions: Takeoff timer displays yellow NO GO.

Procedures:

� Simultaneously reduce collective and establish a decelerative attitude.

� When approaching the ground, establish a landing attitude.

� Apply power to cushion landing.

HIGE AT 3.5 FT (1M)SKID HEIGHT

TDP

8 SECONDS

REJECTED TAKEOFF DISTANCE F927−042A

Figure 11−3.3. Category A Rejected Takeoff





ENGINE FAILURE AFTER TDP

Indications: Normal engine out indications.

Conditions: Category A timer displays GO.

Procedures:

� Decrease collective to prevent rotor droop and adjust power to OEI 2.5 minutelimit.

� Continue takeoff/climb and accelerate to above 40 KIAS (74 Km/H).

� Assure power set to OEI 2.5 minute limits.

� Climb at 45 KIAS (83 Km/H) to 200 FT (61M) HAT.

� Accelerate to VY .

� Climb at VY and OEI MCP (Ref. Section II).

� Refer to Section III for single engine emergencies.

� Takeoff timer − OFF

2−1/2MIN. OEI LIMIT

HIGE AT 3.5 FT (1M)SKID HEIGHT

TDP

CONTINUED TAKEOFF DISTANCE

ACCELERATE TO VTOSS (40 KIAS)

(74KM/H)

CLIMB OEI MCP

AT VY

35 FT

8 SECONDS

200 FT (61M) HAT

ACCELERATE TO VY

F927−043A

CLIMB AT 45 KIAS(83KM/H)

(11m)

Figure 11−3.4. Continued Takeoff





11−3.2. HELIPORT/ELEVATED HELIPAD TAKEOFF PROCEDURES

Normal heliport takeoff and takeoff path:

NOTE: TDP is a point 100 FT HAT and approximately 300 FT behind center of heliport.The distance behind the helipad is achieved by following the takeoff procedurebelow.

� Power assurance check PASS� Pre takeoff check PERFORM� Heliport elevation NOTE HELIPORT ELEVATION WHILE AT FLAT PITCH� Takeoff ESTABLISH CORRECT SIGHT PICTURE BY CLIMBING

VERTICALLY UNTIL THE FAR EDGE OF THEHELIPORT IS JUST ABOVE THE SIGHT PLANE OF THEINSTRUMENT PANEL GLARE SHIELD.CONTINUE REARWARD CLIMB MAINTAINING SAMESIGHT PICTURE TO 130 FT (40M) ABOVE HELIPORTUSING THE BAROMETRIC ALTIMETER (100 FT [30M])HAT).

� At TDP PITCH NOSE DOWN TO TRANSITION TO LEVEL ORCLIMBING FLIGHT WHILE ACCELERATING TO VY

ACCELERATE TO VY

TDP

CLIMB AT VY

ÁÁÁÁÁÁÁÁÁÁÁÁÁÁÁÁ

HELIPAD

F927−044

Figure 11−3.5. Normal Takeoff Profile − Heliport/Elevated Helipad





ENGINE FAILURE PRIOR TO TDP

Indications: Normal engine failure indications.

Conditions: Before reaching or at TDP.

Procedures:

� Pitch nose down (initial pitch down attitude varies with height above the heli-pad starting at 0° at 10 FT (3M) and varying to 20° at TDP) and reduce collectiveto prevent rotor droop.

� When approaching the ground establish a landing attitude. Apply power tocushion landing.

ENGINE FAILURE AFTER TDP


Conditions: After TDP and initiation of forward acceleration (nose down pitch).

Procedures:

� Pitch nose down and reduce collective to prevent rotor droop, accelerate toVTOSS (40 KIAS [74 Km/H]) and adjust power to 2.5 minute OEI limit (Ref.Section II).

NOTE: Initial pitch down attitude is determined by the airspeed at the time of enginefailure − up to 20 degrees nose down prior to an indication of airspeed.

� Continue climb at 45 KIAS (83 Km/H) to 200 FT (61M) HAT.

� Accelerate to VY.

� Continue climb at VY and observe OEI limits (Ref. Section II).






11−3.3. LANDING PROCEDURES − CLEAR AIRFIELD, HELIPORT AND ELEVATED HELIPAD

Normal landing profile:

NOTE: LDP is 100 feet (30 meters) above intended landing area at an airspeed of 35KIAS and a rate of descent of 500 ft/pm (152 m/min) or less.

� Before landing checks PERFORM� Approach angle ESTABLISH A 6° SIGHT PICTURE AND PLAN

APPROACH TO ARRIVE AT LDP AT 35 KIAS� Landing TERMINATE APPROACH ABOVE LANDING AREA

AT 3.5 FT (1M) SKID HEIGHT

HOVER AT 3.5 FT (1M)SKID HEIGHT

LDP 100 FT (30M) AGLAND 35 KIAS

(65KM/H)

F927−045A

Figure 11−3.6. Normal Landing Profile





ENGINE FAILURE PRIOR TO LDP

Indications: Normal engine failure indications

Conditions: Prior to LDP

Procedures:

NOTE: The pilot may elect to perform the following procedures or continue the approachand landing by following the procedures stated for ‘‘Engine Failure After LDP’’.

� Increase power to OEI 2.5 minute limit.

� Accelerate to above 40 KIAS (74 Km/H).

� Climb to 200 FT (61m) at 45 KIAS (83 Km/H).

� Accelerate to VY.

� Continue climb at VY and observe OEI limits (Ref. Section II).



LDP35 KIAS (65KM/H)100 FT (30M) AGL

ENGINE FAILURE PRIOR TO LDP

CLIMB OEI MCPAT VY

OEI MCP

ACCELERATE TO VY

F927−046A

CLIMB AT 45 KIAS(83KM/H)

200 FT (61M) HAT

Figure 11−3.7. Balked Landing





ENGINE FAILURE AFTER LDP


Conditions: After to LDP

Procedures:

� Continue approach.

� Perform OEI landing.


50/25 FT

ENGINE FAILUREAFTER LDP

DECELERATE

CATEGORY A LANDING DISTANCEOPEN AIRFIELD/HELIPORT/ELEVATED HELIPAD F927−047A

LDP35 KIAS (65KM/H)100 FT (30M) AGL

(15/8m)

Figure 11−3.8. Continued Landing





11−3.4. EQUIPMENT MALFUNCTIONS

IIDS FAILURE

Indications: IIDS displays blanks.

Conditions: Loss of electrical power to IIDS.

Procedures: On ground

� Shut down

Procedures: In flight

� Reduce airspeed to 75 KIAS (139 Km/H).

� Reduce electrical load.


Category A OperationsPerformance Data





11−5.1. TAKEOFF PERFORMANCE

Takeoff performance:

Takeoff performance is limited by weight/altitude/OAT limits (Ref. Figure 11−2.2and Figure 11−2.3).

11−5.2. TAKEOFF DISTANCE REQUIRED

Description: Flight planning must be based on the rejected and continued takeoffdistance charts (Ref. Figure 11−5.1, thru Figure 11−5.3.) and the respective chartsfor Segment I and Segment II climb gradients and acceleration distance (Ref.Figure 11−5.5 thru Figure 11−5.7)


Example:Wanted: Takeoff distance required.Known: Maximum takeoff gross weight from example in paragraph

11−2.2 = 6160 LB (2791KG).Known: Outside air temperature = 28°CKnown: Pressure altitude = 2000 FT (610M)Method: Enter top chart at 28°C, move right to the 2000 FT (610M) pressure

altitude line, now move down to the 6160 LB (2791KG) weightpoint (interpolated) and now to the left to read approximately 750 FT (229M) takeoff distance.

NOTE: Rejected Takeoff Distance Required (Clear Airfield) for the above example isapproximately 680 FT (207M) (Ref. Figure 11−5.2)

NOTE: Distance Required to Clear a 35 FT Obstacle for the above example isapproximately 370 FT (113M) (Ref. Figure 11−5.3)





F927−036B

300

500

700

900

1100

1300

1500

1700

1900

0

10

20

30

40

50

−10

−20

SL

TAKEOFF DISTANCEFEET (METERS)

OAT − °C



(METERS)

(KILOGRAMS)6500

5500

6000

5000(2268kg)

(2495kg)

(2722kg)

(2948kg)

−1000(−305M)

10000 HD

LIMIT

(3048M)

1000(305M)

2000(610M)

3000(914M) 4000

(1219M)

5000(1524M)

6000(1829M)

7000(2134M)

10000(3048M)

8000(2438M)

9000(2743M)

(91M)

(152M)

(213M)

(274M)

(335M)

(396M)

(457M)

(518M)

(579M)

Figure 11−5.1. Distance Required to Clear a 35 FT (11M) Obstacle on Takeoff (Clear Airfield)





F927−050B

0

10

20

30

40

50

−10

−20

TAKEOFF DISTANCE

OAT − °C


SL

FEET (METERS)

6250

6500

300

400

500

600

700

800

900

1000

1200

1300

1400

1500

1600

1100

(KILOGRAMS)

−1000(−305M)

11000(3353M)

10000 HD

LIMIT

(3048M)

5000(2268kg)

5200(2359kg)

5400(2449kg)5600

(2540kg)5800

(2631kg)6000

(2722kg)(2835kg)

(2948kg)


(METERS)

(91m)

(122m)

(152m)

(183m)

(213m)

(238M)

(274m)

(305m)

(335m)

(366m)

(396m)

(427m)

(457m)

(488m)

6000(1829M)

5000(1524M)

3000(914M)

2000(610M)

1000(305M)

4000(1219M)

7000(2134M)

10000(3048M)

8000(2438M)

9000(2743M)

Figure 11−5.2. Rejected Takeoff Distance Required (Clear Airfield)





F927−051B

−6

,00

0

7,000

8,000

9,000

10,0

00

11,0

00

TAKEOFF DISTANCE

OAT − °C


0

10

20

30

40

50

−10

−20

SL

6500

5500

6000

5000

0

100

200

300

400

500

600

700

(91m)

(122m)

(152m)

(183m)

(213m)

(30M)

(61M)

(2268kg)(2495kg)

(2722kg)

(2948kg)

10000 HD

LIMIT

(3048M)


(METERS)

(KILOGRAMS)

−1000(−305M)

6000(1829M)

5000(1524M)

7000(2134M)

10000(3048M)

8000(2438M)

9000(2743M)

3000(914M)

2000(610M)

1000(305M)

4000(1219M)

FEET (METERS)

Figure 11−5.3. Distance Required to Clear a 35 FT (11M)Obstacle on Takeoff Heliport/Elevated Helipad





11−5.3. CONTINUED TAKEOFF FLIGHT PATH

The continued takeoff flight path begins at the end of the Continued Takeoff DistanceRequired, at 35 feet (11 meters) above the takeoff surface or higher at VTOSS, andis divided into three segments.

SEGMENT ICLIMB AT 45 KIAS

(83KM/H)

F927−048A


SEGMENT IICLIMB AT VY

ACCELERATE TO VY

200 FT (61M) AGL

CONTINUOUS OEI LIMIT

1000 FT(305M) AGL

Figure 11−5.4. OEI Takeoff Flight Path





F927−037B

OAT − °C

0

10

20

30

40

50

−10

−20

SL

11000

500

1000

1500

2000

2500

3000

3500

4000



(METERS)

6000(1829M)

5000(1524M)3000

(914M)

2000(610M)

1000(305M)

4000(1219M)

−1000(−305M)

7000(2134M)

10000(3048M)

8000(2438M)

9000(2743M)

(3353M)

10000 HD

LIMIT

(3048M)

GROSS WEIGHT − POUNDS(KILOGRAMS)

62506500

5000(2268kg)

5200(2359kg)

5400(2449kg)

5600(2540kg)

5800(2631kg)6000

(2722kg)(2835kg)

(2948kg)

(152M)

(305M)

(457M)

(610M)

(762M)

(914M)

(1067M)

(1219M)

Figure 11−5.5. Takeoff Distance Segment I − Distance Required to Climb from 35 FT(11M) to 200 FT (61M) HAT





F927−038B

OAT − °C

0

10

20

30

40

50

−10

−20

SL

2100

PRESSURE ALTITUDE

X 1000 FT

TAKEOFF DISTANCE − FEET(METERS)

10000 HD

LIMIT

(3048M)11000

6000(1829M)

5000(1524M)

3000(914M)

2000(610M)

1000(305M)

4000(1219M)

−1000(−305M)

7000(2134M)

10000(3048M)

8000(2438M)

9000(2743M)

(3353M)


62506500

5000(2268kg)

5200(2359kg)

5400(2449kg)

5600(2540kg)

5800(2631kg)

(2835kg)(2948kg)

6000(2722kg)

500

700

900

1100

1300

1500

1700

1900

(152M)

(213M)

(274M)

(335M)

(396M)

(457M)

(518M)

(579M)

(640M)

Figure 11−5.6. Acceleration Distance − Distance to Accelerate from 45 KIAS (83 KM/H) to VY at 200 FT (61M) HAT





5000(1524M)

F927−039B

OAT − °C

0

10

20

30

40

50

−10

−20

5000

10000

15000

20000

25000

30000

35000

400005000 (2268kg)

5200 (2359kg)

5400 (2449kg)

5600 (2540kg)

5800 (2631kg)

6000 (2722kg)

6250 (2835kg)

6500 (2948kg)


(METERS)

11000

6000(1829M)

3000(914M)1000

(305M)

4000(1219M)

7000(2134M)

10000(3048M)

8000(2438M)

9000(2743M)

(3353M)

−1000(−305M)

2000(610M)


(1524m)

(3048m)

(4572m)

(6096m)

(7620m)

(9144m)

(10668m)

(12192m)


10000 HD

LIMIT

(3048M)

Figure 11−5.7. Takeoff Distance Segment II − Distance Required to Climb from 200 FT (61M) HAT to 1000 FT (305M)





11−5.4. LANDING PERFORMANCE − OPEN AIRFIELD

The landing distance from 50 FT (15m) above the landing surface to the point atwhich the helicopter comes to a complete stop is 500 FT (152m).

11−5.5. LANDING PERFORMANCE − HELIPORT/ELEVATED HELIPAD

The landing distance from 25 FT (8m) above the landing surface to the point atwhich the helicopter comes to a complete stop is 250 FT (76m).

Category A OperationsAdditional Operations




PART IXADDITIONAL OPERATIONS

11−9.1. CATEGORY A OEI TRAINING

To simulate OEI operations in a Category A environment, follow the profiles asdescribed in ‘‘Part 3, Takeoff and Landing Procedures’’, of this section.

Operate at the recommended gross weight as depicted in Figure 11−9.1. See examplebelow.

Observe normal (twin) engine operating limitations (Ref. Section II).

NOTE: Operating at the recommended gross weight, assists the pilot in maintainingnormal (twin) engine operating limitations and accurately simulates actual OEIconditions.

Description: This chart (Ref. Figure 11−9.1) reflects the weight at which CategoryA OEI training may be performed with the operating engine within normal (twin)engine operating limitations.


Example:

Wanted: Maximum gross weight for training under the following conditions.Known: HP = 4000 FT (1219m), OAT = 10°C

Method: Enter the chart at 10°C and move vertically to 4000 HP (1219m) curve.At this point move directly to the left and read from the gross weight scale,5200 LB (2359kg).

Category A OperationsAdditional Operations




F927−049B

−20 −10 0 10 20 30 40 50

3700

3800

3900

4000

4100

4200

4300

4400

4500

4600

4700

4800

4900

OAT − °C−36 −30

SL

PR

ES

SU

RE

ALT

ITU

DE

− F

EE

T (

M)

GR

OS

S W

EIG

HT

− P

OU

ND

S (

KG

)

7000(2134M)

6000(1829M)

5000(1524M)

4000(1219M)

3000(914M)

2000(610M)

1000(305M)

−1000

(−305M)

10000 HD

LIMIT

(3048M)

11000(3353M)

10000(3048M)

8000(2438M)

9000(2743M)

13000(3962M)

12000(3658M)

14000(4267M)

5000

5100

5200

5300

5400

5500

5600

5700

(2268kg)

(2313kg)

(2359kg)

(2404kg)

(2449kg)

(2495kg)

(2540kg)

(2585kg)

(2223kg)

(2177kg)

(2132kg)

(2087kg)

(2041kg)

(1996kg)

(1950kg)

(1905kg)

(1860kg)

(1814kg)

(1769kg)

(1724kg)

(1678kg)

Figure 11−9.1. Recommended OEI Training Weight

Rotorcraft Flight Manual

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