WORKBOOK - Naval Air Training Command

NAVAL AIR TRAINING COMMAND

NAS CORPUS CHRISTI, TEXAS CNATRA P-421 (Rev. 03-21)

WORKBOOK

INTRODUCTION TO

HELICOPTER AERODYNAMICS

TH-57

2021

DEPARTMENT OF THE NAVY CHIEF OF NAVAL AIR TRAINING250 LEXINGTON BLVD SUITE 102 CORPUS CHRISTI TX 78419-5041

CNATRA P-421 N714 19 Mar 21

CNATRA P-421 (REV. 3-21)

Subj: WORKBOOK, INTRODUCTION TO HELICOPTER AERODYNAMICS, TH-57

1. CNATRA P-421 (Rev. 3-21) PAT, “Workbook, Introduction to Helicopter Aerodynamics, TH-57” is issued for information, standardization of instruction, and guidance to all flight instructors and student military aviators within the Naval Air Training Command.

2. This publication is an explanatory aid to the Helicopter and Tiltrotor curriculums and shall be the authority for the execution of all flight procedures and maneuvers herein contained.

3. Recommendations for changes shall be submitted via the electronic Training Change Request (TCR) form located on the CNATRA Website.

4. CNATRA P-421 (New 02-21) PAT is hereby cancelled and superseded.

D. F. WESTPHALLBy direction

Releasability and distribution: This instruction is cleared for public release and is available electronically only via Chief of Naval Air Training Issuances Website, https://www.cnatra.navy.mil/pubs-pat-pubs.asp.

https://www.cnatra.navy.mil/pubs-pat-pubs.asp

iii

INTRODUCTION TO HELICOPTER

AERODYNAMICS WORKBOOK

TH-57

Q-2C-3156

iv

LIST OF EFFECTIVE PAGES

Dates of issue for original and changed pages are:

Original.. .0...18 Feb 21 (this will be the date issued)

Revision...1...19 Mar 21

TOTAL NUMBER OF PAGES IN THIS PUBLICATION IS 178 CONSISTING OF THE FOLLOWING:

Page No. Change No. Page No. Change No.

COVER 0 A-1 – A-16 0

LETTER 0 B-1 – B-2 0

iii – xiv 0

1-1 – 1-8 0

2-1 – 2-12 0

3-1 – 3-17 0

3-18 (blank) 0

4-1 – 4-11 0

4-12 (blank) 0

5-1 – 5-18 0

6-1 – 6-16 0

7-1 – 7-28 0

8-1 – 8-19 0

8-20 (blank) 0

9-1 – 9-13 0

9-14 (blank) 0

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INTERIM CHANGE SUMMARY

The following Changes have been previously incorporated in this manual:

CHANGE

NUMBER REMARKS/PURPOSE

The following Interim Changes have been incorporated in this Change/Revision:

INTERIM

CHANGE

NUMBER

REMARKS/PURPOSE

ENTERED

BY

DATE

vi

ACKNOWLEDGEMENTS

The genesis for this text was direction from CNATRA to develop a more comprehensive

helicopter aerodynamics text and course with the ultimate goal of making smarter and therefore

safer pilots. This text was produced through the collective effort of Training Air Wing FIVE

academic instructors, helicopter pilots, the Rotary Wing Aerodynamics Instructor from the U.S.

Navy & Marine Corps School of Aviation Safety, and professors from the U.S. Naval Academy.

The text is a reference that the fleet pilot can use as a single source document, not just as a flight

school workbook. It includes a great deal of information and pictorial items from open internet

sources as well as the books in the reference list. If there are any errors in the interpretation of

these authors’ materials or the sources in the reference list (Appendix A), or the information as

presented, the reader should report them to Training Air Wing FIVE Academic Training

Department. This book is a work in progress and will be updated periodically.

vii

INSTRUCTIONS FOR STUDENT NAVAL AVIATORS

Objective

Upon completion of this course, the student will possess an understanding of basic helicopter

fundamentals and aerodynamic principles. While the student will be required to demonstrate a

functional knowledge of the material presented through successful completion of an end-of-

course examination with a minimum score of 80%, this course is primarily focused on preparing

the student for practical application of that knowledge in future helicopter flight and flight

planning.

Assumptions

1. An engineering background is not required for mastery of the basic concepts of this course.

2. Recent completion of the Fundamentals of Aerodynamics course offered in Aviation

Preflight Indoctrination (or an equivalent course). Review the material if an extended period of

time has passed or the material was not previously mastered.

3. Concurrent completion of interactive courseware (ICWs) augmented by the availability of

knowledgeable instructors and a formal review session prior to testing.

4. This document is an introductory text that serves as an adequate, stand-alone, ready

reference for the military helicopter pilots. A comprehensive helicopter engineering text is not

intended or necessary. A recommended reading list provides options for those with greater

curiosity.

Instructional Approach of this edition

1. This edition was reorganized to be more user friendly. The concepts to be mastered for the

end-of-course exam are described in the corresponding enabling objectives in the basic text.

2. This edition includes the Table of Contents and appendices with a reference/recommended

reading list, Glossary and an Index.

Recommendations for Students

1. Review the enabling objectives before completing ICW’s, study the text and the class

presentation.

2. The utility of this course depends primarily upon the conscientious accomplishment of your

reading and study assignments.

3. Participation in a study group is highly recommended. A study group of four is optimum.

viii

AVIATOR AND INSTRUCTOR GUIDANCE

Every aviator has a “toolbox” with tools they have collected and mastered over the years.

Among other things, that toolbox probably contains emergency procedures, operating

limitations, regulations, or instructions, and an assortment of lessons learned.

A mastery of basic aerodynamic principles is an important tool for the professional aviator,

especially when it comes to the challenging rigors of rotary wing flight in military aviation.

Aerodynamics can save your life. Time after time, mishap reports attribute either a lack of

understanding or the inappropriate misapplication of aerodynamic principles as a causal factor of

the mishap.

The exposure in flight training to these principles and their application is only the first step in the

mastery of the essential concepts. Every profession requires continuing education. As iron

sharpens iron, we learn from one another. It is the responsibility of every naval aviator to

continually sharpen the skills necessary to maintain the “edge” that may one day make the

difference in the accomplishment of the mission and/or your crew’s survival.

Periodic review of both systems and aerodynamic course material will provide for the greatest

retention and immediate recall. Each time an aviator re-reads the text, they will glean some new

fact or relationship to improve their overall understanding, including those who have yet to set

foot in a new aircraft.

A survey of pilots of the Aviation Safety Officer Course reveals that few of the aerodynamic

principles necessary to investigate a mishap were retained from flight training. Each ASO

completes twenty lecture hours of rotary wing aerodynamics in the course. It is clear from this

course that every individual is not only capable of, but also highly motivated toward, making

basic aerodynamic principles an integral part of their toolbox. Their goal is not to develop

mishap investigation skills, but rather to develop mishap avoidance skills they can share.

Therefore, the goal of this text is to present the principles of helicopter aerodynamics in a

straightforward, comprehensible manner, such that both the newest student naval aviator and the

crustiest old instructor pilot may have at their disposal a concise, accurate reference. The best

available tool, however, is only of use to the craftsman who develops and maintains a level of

expertise to make its use second nature.

ix

TABLE OF CONTENTS

LIST OF EFFECTIVE PAGES .................................................................................................. iv INTERIM CHANGE SUMMARY ...............................................................................................v ACKNOWLEDGEMENTS ........................................................................................................ vi INSTRUCTIONS FOR STUDENT NAVAL AVIATORS ..................................................... vii

AVIATOR AND INSTRUCTOR GUIDANCE....................................................................... viii TABLE OF CONTENTS ............................................................................................................ ix TABLE OF FIGURES ................................................................................................................ xii

CHAPTER ONE - INTRODUCTION TO THE HELICOPTER ......................................... 1-1 100. INTRODUCTION .................................................................................................. 1-1

101. HELICOPTER AERODYNAMICS COURSE LEARNING OBJECTIVES ........ 1-1

102. ROTOR SYSTEM .................................................................................................. 1-2 103. ROTOR CONFIGURATIONS ............................................................................... 1-3

104. CONTROLLING FLIGHT ..................................................................................... 1-6

105. FLIGHT CONDITIONS ......................................................................................... 1-7

CHAPTER TWO - THE ATMOSPHERE .............................................................................. 2-1 200. INTRODUCTION .................................................................................................. 2-1 201. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 2-1

202. REVIEW OF BASIC PHYSICS AND AERODYNAMICS .................................. 2-1 203. VECTOR ANALYSIS ............................................................................................ 2-1

204. ESSENTIAL TERMS AND DEFINITIONS ......................................................... 2-3 205. PROPERTIES OF THE ATMOSPHERE .............................................................. 2-4 206. THE STANDARD ATMOSPHERE ...................................................................... 2-5

207. ALTITUDE COMPUTATIONS ............................................................................ 2-6

208. ALTITUDE COMPUTATIONS – DENSITY ALTITUDE................................... 2-8

CHAPTER THREE - HELICOPTER AERODYNAMICS BASICS ................................... 3-1 300. INTRODUCTION .................................................................................................. 3-1 301. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 3-1

302. AIRCRAFT REFERENCE SYSTEM .................................................................... 3-1 303. GENERAL .............................................................................................................. 3-2 304. FORCES ACTING ON THE AIRCRAFT ............................................................. 3-4 305. LIFT ........................................................................................................................ 3-4 306. WEIGHT ................................................................................................................. 3-8

307. THRUST ................................................................................................................. 3-9

308. DRAG ..................................................................................................................... 3-9

309. INTRODUCTION TO LIFT THEORIES ............................................................ 3-12 310. PRESSURE DISTRIBUTION THEORY............................................................. 3-12 311. CIRCULATION THEORY .................................................................................. 3-13 312. MOMENTUM THEORY ..................................................................................... 3-14 313. BLADE ELEMENT THEORY ............................................................................ 3-15

x

CHAPTER FOUR - AIRFOILS ............................................................................................... 4-1 400. INTRODUCTION .................................................................................................. 4-1 401. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 4-1

402. AIRFOIL ................................................................................................................. 4-1 403. AIRFOIL TERMINOLOGY AND DEFINITIONS ............................................... 4-1 404. AIRFOIL TYPES.................................................................................................... 4-3 405. ROTOR AXIS ......................................................................................................... 4-4 406. AIRFLOW AND REACTIONS IN THE ROTOR DISK ...................................... 4-4

407. ROTOR BLADE ANGLES .................................................................................... 4-9

CHAPTER FIVE - POWERED FLIGHT ............................................................................... 5-1 500. INTRODUCTION .................................................................................................. 5-1 501. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 5-1

502. POWERED FLIGHT .............................................................................................. 5-1 503. HOVERING FLIGHT ............................................................................................ 5-1

504. VERTICAL FLIGHT.............................................................................................. 5-7 505. FORWARD FLIGHT ............................................................................................. 5-8

506. SIDEWARD FLIGHT .......................................................................................... 5-16 507. REARWARD FLIGHT ........................................................................................ 5-17

508. TURNING FLIGHT ............................................................................................. 5-18

CHAPTER SIX - AUTOROTATION ...................................................................................... 6-1 600. INTRODUCTION .................................................................................................. 6-1 601. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 6-1

602. GENERAL .............................................................................................................. 6-1 603. VERTICAL AUTOROTATION ............................................................................ 6-2

604. FORWARD FLIGHT AUTOROTATION ............................................................. 6-3 605. AUTOROTATION DESCENT VARIABLES....................................................... 6-5

606. PHASES OF AUTOROTATION ........................................................................... 6-6 607. WINDMILL BRAKE STATE .............................................................................. 6-13 608. HEIGHT-VELOCITY DIAGRAM ...................................................................... 6-13

CHAPTER SEVEN - PERFORMANCE ................................................................................. 7-1 700. INTRODUCTION .................................................................................................. 7-1 701. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 7-1 702. FACTORS AFFECTING PERFORMANCE ......................................................... 7-1 703. GENERAL .............................................................................................................. 7-2 704. POWER REQUIRED ............................................................................................. 7-4

705. POWER AVAILABLE ........................................................................................... 7-8 706. EFFECT OF TAIL ROTOR ON POWER AVAILABLE ...................................... 7-9

707. POWER REQUIRED EXCEEDS POWER AVAILABLE ................................. 7-10 708. HOVER PERFORMANCE .................................................................................. 7-12 709. CLIMB PERFORMANCE ................................................................................... 7-17 710. REVIEW OF OPTIMUM AIRSPEEDS ............................................................... 7-19

xi

CHAPTER EIGHT - FLIGHT PHENOMENA...................................................................... 8-1 800. INTRODUCTION .................................................................................................. 8-1 801. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 8-1

802. GENERAL .............................................................................................................. 8-1 803. FLIGHT ENVELOPE / V-N DIAGRAM .............................................................. 8-1 804. VIBRATION ANALYSIS ...................................................................................... 8-5 805. GROUND VORTEX .............................................................................................. 8-6 806. COMPRESSIBILITY ............................................................................................. 8-7

807. RETREATING BLADE STALL .......................................................................... 8-10 808. GROUND RESONANCE .................................................................................... 8-11 809. DYNAMIC ROLLOVER ..................................................................................... 8-11 810. LOW-G CONDITIONS ........................................................................................ 8-16 811. LOW ROTOR RPM AND ROTOR STALL ........................................................ 8-17

CHAPTER NINE - TAIL ROTOR CONSIDERATIONS ..................................................... 9-1 900. INTRODUCTION .................................................................................................. 9-1 901. LESSON TOPIC LEARNING OBJECTIVES ....................................................... 9-1

902. TORQUE EFFECT ................................................................................................. 9-1 903. VERTICAL STABILIZER ..................................................................................... 9-2

904. TRANSLATING TENDENCY AND HOVER ATTITUDE ................................. 9-3 905. WEATHER VANING ............................................................................................ 9-4 906. TAIL ROTOR FAILURES AND ISSUES ............................................................. 9-4

APPENDIX A - GLOSSARY ................................................................................................... A-1 A100. GLOSSARY .......................................................................................................... A-1

APPENDIX B - REFERENCES AND READING LIST .......................................................B-1

xii

TABLE OF FIGURES

Figure 1-1 Basic Components of the rotor system ............................................................. 1-3 Figure 1-2 Tandem Rotor Helicopters ................................................................................ 1-4 Figure 1-3 Coaxial rotors ..................................................................................................... 1-4 Figure 1-4 HH-43 Hiskie with intermeshing rotors ........................................................... 1-5 Figure 1-5 Basic tail rotor components............................................................................... 1-6

Figure 2-1 Resultant by the Tip-to-Tail Method ............................................................... 2-2 Figure 2-2 Force Vectors on an Airfoil ............................................................................... 2-2 Figure 2-3 Force Vectors on Aircraft in Flight .................................................................. 2-2 Figure 2-4 Static pressure .................................................................................................... 2-4 Figure 2-5 Molecular Energy and Air Temperature ......................................................... 2-5

Figure 2-6 Standard Atmospheric Table ............................................................................ 2-6 Figure 2-7 Pressure Altitude vs. True Altitude with Varying Local Pressures .............. 2-7

Figure 2-8 DA Chart............................................................................................................. 2-9

Figure 2-9 Dew Point Correction Chart ........................................................................... 2-10 Figure 2-10 Thrust Variation with Humidity .................................................................... 2-11 Figure 2-11 Sample Effects of DA Calculations on CH-53D Performance ..................... 2-12

Figure 3-1 Aircraft Reference System ................................................................................ 3-2 Figure 3-2 Area of a Blade ................................................................................................... 3-3

Figure 3-3 Profile of an airfoil ............................................................................................. 3-3 Figure 3-4 Forces acting on a helicopter in forward flight ............................................... 3-4

Figure 3-5 Production of lift ................................................................................................ 3-5 Figure 3-6 Water flow through a tube ................................................................................ 3-6

Figure 3-7 Venturi effect ...................................................................................................... 3-7 Figure 3-8 Notional load factor diagram ............................................................................ 3-9 Figure 3-9 Notional Drag profiles ..................................................................................... 3-10

Figure 3-10 Induced Drag .................................................................................................... 3-11 Figure 3-11 Drag Curve ....................................................................................................... 3-12

Figure 3-12 Pressure Changes Around a Cambered Airfoil ............................................ 3-13 Figure 3-13 Magnus Effect................................................................................................... 3-14

Figure 3-14 Pressure Distribution for Magnus Effect ....................................................... 3-14 Figure 3-15 Induced Velocity Idealized for Momentum Theory ..................................... 3-15 Figure 3-16 Variables in the Blade Element Theory ......................................................... 3-16 Figure 3-17 Airflow over the Airfoil ................................................................................... 3-16 Figure 3-18 Blade Element Diagram .................................................................................. 3-17

Figure 4-1 Aerodynamic terms of an airfoil ....................................................................... 4-2

Figure 4-2 Airfoil Types ....................................................................................................... 4-3 Figure 4-3 Blade Twist ......................................................................................................... 4-4 Figure 4-4 Relative Wind ..................................................................................................... 4-5 Figure 4-5 Horizontal component of relative wind............................................................ 4-5 Figure 4-6 Induced Flow ...................................................................................................... 4-5 Figure 4-7 Spanwise induced flow velocities ...................................................................... 4-6

xiii

Figure 4-8 Rotational Relative Wind .................................................................................. 4-6

Figure 4-9 Resultant relative wind ...................................................................................... 4-7

Figure 4-10 Induced flow in forward flight .......................................................................... 4-7 Figure 4-11 In Ground Effect (IGE) ..................................................................................... 4-8 Figure 4-12 Out of Ground Effect (OGE) ............................................................................ 4-9 Figure 4-13 Angle of Incidence .............................................................................................. 4-9 Figure 4-14 Angle of Attack ................................................................................................. 4-10

Figure 5-1 Translating Tendency ........................................................................................ 5-3 Figure 5-2 Pendular action .................................................................................................. 5-4 Figure 5-3 Coning ................................................................................................................. 5-5 Figure 5-4 Gyroscopic precession ....................................................................................... 5-6 Figure 5-5 Forward Cyclic Input ........................................................................................ 5-7

Figure 5-6 No wind hover .................................................................................................... 5-8 Figure 5-7 Transition to forward flight .............................................................................. 5-8

Figure 5-8 Power vs. airspeed chart.................................................................................... 5-9 Figure 5-9 Airflow in forward flight ................................................................................. 5-10

Figure 5-10 Dissymmetry of lift ........................................................................................... 5-11 Figure 5-11 Effect of flapping .............................................................................................. 5-12

Figure 5-12 Blowback ........................................................................................................... 5-13 Figure 5-13 Airflow with minimal headwind ..................................................................... 5-14 Figure 5-14 Airflow just prior to ETL ................................................................................ 5-14

Figure 5-15 ETL.................................................................................................................... 5-15 Figure 5-16 Sideward Flight ................................................................................................ 5-16

Figure 5-17 Rearward Flight ............................................................................................... 5-17 Figure 5-18 Turning Flight .................................................................................................. 5-18

Figure 6-1 Airflow in an autorotation................................................................................. 6-2 Figure 6-2 Rotor disc regions in autorotation zero speed ................................................. 6-2

Figure 6-3 Rotor blade regions in autorotation ................................................................. 6-4 Figure 6-4 Rotor disc regions in autorotation forward speed .......................................... 6-5

Figure 6-5 Force Vectors in Level-Powered Flight at High Speed................................... 6-6 Figure 6-6 Force Vectors after Power Loss – Reduced Collective ................................... 6-7

Figure 6-7 Force Vectors in Autorotative Steady-State Descent ...................................... 6-8 Figure 6-8 RPM Response to Small RPM Variations ....................................................... 6-9 Figure 6-9 Autorotational Rate of Descent Compared to Airspeed ............................... 6-11 Figure 6-10 Blade Element and Thrust during Steady State Auto and Flare................. 6-12 Figure 6-11 Generic Height Velocity Diagram .................................................................. 6-14

Figure 7-1 Total Drag Curve ............................................................................................... 7-3

Figure 7-2 Power Required Versus Airspeed Curve ......................................................... 7-4 Figure 7-3 Aerodynamic Forces Affecting Power Required ............................................ 7-5 Figure 7-4 Power Required Curves versus Airspeed ........................................................ 7-5 Figure 7-5 Optimum Airspeeds ........................................................................................... 7-7 Figure 7-6 Maximum Range Airspeed Adjustment for Winds ........................................ 7-8 Figure 7-7 Induced Power Required ................................................................................. 7-11

xiv

Figure 7-8 Parasite Power Required ................................................................................. 7-11

Figure 7-9 Decrease in Excess Power as Airspeed Decreases ......................................... 7-12

Figure 7-10 Sample hover problem 1 .................................................................................. 7-14 Figure 7-11 Sample hover problem 2 .................................................................................. 7-15 Figure 7-12 Sample hover problem 3 .................................................................................. 7-16 Figure 7-13 Sample cruise problem .................................................................................... 7-18 Figure 7-14 Sample climb problem ..................................................................................... 7-19 Figure 7-15 Optimum Airspeeds ......................................................................................... 7-20

Figure 7-16 Power Required Chart (CH-46E) ................................................................... 7-21 Figure 7-17 Fuel Flow vs. TAS ............................................................................................ 7-22 Figure 7-18 Excess Power .................................................................................................... 7-22 Figure 7-19 Max Range Altitude vs. Gross Weight ........................................................... 7-23 Figure 7-20 AH-1W - Max Range Airspeed vs. Gross Weight ......................................... 7-24

Figure 7-21 MH-60S – Max Range Airspeed vs. Gross Weight ....................................... 7-25 Figure 7-22 RPM vs. Fuel Flow ........................................................................................... 7-26 Figure 7-23 Excess Power .................................................................................................... 7-27 Figure 7-24 Best Angle of Climb ......................................................................................... 7-28

Figure 7-25 Rate of Climb vs. Best Angle of Climb ........................................................... 7-28

Figure 8-1 V-n Diagram for Fixed Wing Aircraft ............................................................. 8-2 Figure 8-2 AH-64 Apache V-n Diagram ............................................................................. 8-3 Figure 8-3 Vibration Analysis ............................................................................................. 8-6

Figure 8-4 Ground Vortex ................................................................................................... 8-7 Figure 8-5 Critical Mach Number ...................................................................................... 8-9

Figure 8-6 Drag Increase with Mach Number ................................................................... 8-9 Figure 8-7 Dynamic rollover .............................................................................................. 8-13 Figure 8-8 Slope takeoff or landing 1................................................................................ 8-14 Figure 8-9 Slope takeoff or landing 2................................................................................ 8-15

Figure 9-1 Tail Rotor Unbalanced Force ........................................................................... 9-2 Figure 9-2 Yaw Control Mechanisms for Various Configurations .................................. 9-2

Figure 9-3 Vertical Stabilizer .............................................................................................. 9-3 Figure 9-4 Translating Tendency ........................................................................................ 9-4

Figure 9-5 Weather Vaning ................................................................................................. 9-4 Figure 9-6 Effects of Wind Direction on Directional Control .......................................... 9-9 Figure 9-7 LTE in a Right Crosswind .............................................................................. 9-11 Figure 9-8 Fly Home Capability After Loss Of Tail Rotor Thrust ................................ 9-13

INTRODUCTION TO THE HELICOPTER 1-1

CHAPTER ONE

INTRODUCTION TO THE HELICOPTER

100. INTRODUCTION

A helicopter is an aircraft lifted and propelled by one or more horizontal rotors. Each rotor consists

of two or more rotor blades. Helicopters are classified as rotorcraft or rotary-wing aircraft to

distinguish them from fixed-wing aircraft because the helicopter derives its source of lift from the

rotor blades rotating around a mast. The word “helicopter” is adapted from the French hélicoptère,

coined by Gustave de Ponton d’Amécourt in 1861, and linked to the Greek words helix/helikos

(“spiral” or “turning”) and pteron (“wing”).

As an aircraft, the primary advantages of the helicopter are due to the rotor blades that revolve

through the air, providing lift without requiring the aircraft to move forward. This lift allows the

helicopter to hover in one area and to take off and land vertically without the need for runways.

For this reason, helicopters are often used in congested or isolated areas where fixed-wing aircraft

are not able to take off or land.

Piloting a helicopter requires adequate, focused, and safety-orientated training. It also requires

continuous attention to the machine and the operating environment. The pilot must work in three

dimensions and use both arms and both legs constantly to keep the helicopter in a desired state.

Coordination, timing, and control touch require simultaneous use when flying a helicopter.

Although helicopters were developed and built during the first half-century of flight, some even

reaching limited production; it was not until 1942 that a helicopter designed by Igor Sikorsky

reached full-scale production, with 131 aircraft built. Even though most previous designs used

more than one main rotor, it was the single main rotor with an anti-torque tail rotor configuration

that was recognized worldwide as the helicopter.

101. HELICOPTER AERODYNAMICS COURSE LEARNING OBJECTIVES

1. Identify the basic physics principles needed to support helicopter flight

2. Identify the basic aerodynamic factors that are vital to helicopter performance

3. Identify airfoil design considerations

4. Identify the three types of rotor systems

5. Identify rotor system dynamics

6. Identify rotorcraft configurations and airfoil design considerations

7. Identify the basic aerodynamic characteristics of the airframe

8. Identify factors that affect helicopter stability and control

CHAPTER ONE HELICOPTER AERODYNAMICS WORKBOOK

1-2 INTRODUCTION TO THE HELICOPTER

9. Identify factors that affect helicopter power required and power available for flight

10. Identify aerodynamic terms, concepts, and diagrams vital to helicopter aerodynamics

11. Explain the aerodynamics of flight

12. Identify factors that lead to undesirable helicopter phenomena

13. Identify actions that prevent undesirable helicopter phenomena

14. Explain undesirable helicopter phenomena

102. ROTOR SYSTEM

The helicopter rotor system is the rotating part of a helicopter that generates lift. A rotor system

may be mounted horizontally, as main rotors are, providing lift vertically; and it may be mounted

vertically, such as a tail rotor, to provide lift horizontally as thrust to counteract torque effect. In

the case of tilt rotors, the rotor mounts on a nacelle that rotates at the edge of the wing to transition

the rotor from a horizontal mounted position, providing lift horizontally as thrust, to a vertical

mounted position providing lift exactly as a helicopter.

The rotor consists of a mast, hub, and rotor blades. The mast is a hollow cylindrical metal shaft

extending upwards from and driven by the transmission. At the top of the mast is the hub. The

hub is the attachment point for the rotor blades. The rotor blades attach to the hub by several

different methods. Main rotor systems are classified according to how the main rotor blades are

attached and move relative to the main rotor hub. There are three basic classifications: semi rigid,

rigid, or fully articulated, although some modern rotor systems use an engineered combination of

these types.

With a single main rotor helicopter, a torque effect generates as the engine turns the rotor. This

torque causes the body of the helicopter to turn in the opposite direction of the rotor (Newton’s

Third Law: Every action has an equal and opposite reaction. To eliminate this effect, some sort

of anti-torque control must be used with a sufficient margin of power available to allow the

helicopter to maintain its heading and prevent the aircraft from moving unsteadily.

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER ONE


Figure 1-1 Basic Components of the rotor system

103. ROTOR CONFIGURATIONS

Most helicopters have a single main rotor but require a separate rotor to overcome torque, which

is a turning or twisting force. This occurs through a variable pitch, anti-torque rotor or tail rotor.

This is the design that Igor Sikorsky settled on for his VS-300 helicopter. It has become the

recognized convention for helicopter design, although designs do vary. Helicopter main rotor

designs from different manufacturers rotate in one of two different directions (clockwise or

counter-clockwise when viewed from above). This can make it confusing when discussing

aerodynamic effects on the main rotor between different designs, since the effects may manifest

on opposite sides of each aircraft. For clarity, throughout this workbook all examples use a

counter-clockwise rotating (when viewed from above) main rotor system.

1. Tandem Rotor. Tandem rotor (sometimes referred to as dual rotor) helicopters have two

large horizontal rotor assemblies, instead of one main assembly and a smaller tail rotor. Single

rotor helicopters need a tail rotor to neutralize the twisting momentum produced by the single

large rotor. Tandem rotor helicopters, however, use counter-rotating rotors, each canceling out

the other’s torque. Counter-rotating rotor blades will not collide with and destroy each other if

they flex into the other rotor’s pathway. This configuration has the advantage of being able to

hold more weight with shorter blades, since there are two blade sets. This configuration allows

all of the power from the engines available for lift, whereas a single rotor helicopter must use

some power to counter main rotor torque. Because of this, tandem helicopters make up some of

the most powerful and fastest rotor system aircraft.



Figure 1-2 Tandem Rotor Helicopters

2. Coaxial Rotors. Coaxial rotors are a pair of rotors turning in opposite directions, but

mounted on a mast with the same axis of rotation, one above the other. This configuration is a

noted feature of helicopters produced by the Russian Kamov helicopter design bureau.

Figure 1-3 Coaxial rotors

3. Intermeshing Rotors. Intermeshing rotors on a helicopter are a set of two rotors turning

in opposite directions, with each rotor mast mounted on the helicopter with a slight angle to the

other so that the blades intermesh without colliding. This arrangement allows the helicopter to



function without the need for a tail rotor. It has high stability and powerful lifting capability.

This configuration is sometimes referred to as a synchropter. The arrangement was developed in

Germany for a small anti-submarine warfare helicopter, the Flettner Fl 282 Kolibri. During the

Cold War, the American Kaman Aircraft company produced the HH-43 Huskie, for USAF

firefighting purposes. The latest Kaman K-MAX model is a dedicated sky crane design used for

construction work.

Figure 1-4 HH-43 Hiskie with intermeshing rotors

4. Tail Rotor. The tail rotor is a smaller rotor mounted vertically or near-vertically on the tail

of a traditional single-rotor helicopter. The tail rotor either pushes or pulls against the tail to

counter the torque. The tail rotor drive system consists of a drive shaft, powered from the main

transmission, and a gearbox mounted at the end of the tail boom. The drive shaft may consist of

one long shaft or a series of shorter shafts connected at both ends with flexible couplings. The

flexible couplings allow the drive shaft to flex with the tail boom.

The gearbox at the end of the tail boom provides an angled drive for the tail rotor and may

include gearing to adjust the output to the optimum rotational speed typically measured in

revolutions per minute (rpm) for the tail rotor. On some larger helicopters, one or more

intermediate gearboxes angle the tail rotor drive shaft from along the tail boom or tail cone to the

top of the tail rotor pylon. These also serve as a vertical stabilizing airfoil to alleviate the power

requirement for the tail rotor in forward flight. The pylon (or vertical fin) may also provide

limited anti-torque within certain airspeed ranges if the tail rotor or the tail rotor flight controls

fail.



Figure 1-5 Basic tail rotor components

104. CONTROLLING FLIGHT

A helicopter has four primary flight controls: Cyclic, Collective, Anti-torque pedals, and

Throttle

1. Cyclic. The cyclic control is usually located between the pilot’s legs and is commonly

called the “cyclic stick” or simply “cyclic.” On most helicopters, the cyclic is similar to a

joystick. The control is called the cyclic because it can vary the pitch of the rotor blades

throughout each revolution of the main rotor system (i.e., through each cycle of rotation) to

develop unequal lift (thrust). The result is to tilt the rotor disk in a particular direction, resulting

in the helicopter moving in that direction. If the pilot pushes the cyclic forward, the rotor disk

tilts forward, and the rotor produces a thrust in the forward direction. If the pilot pushes the

cyclic to the side, the rotor disk tilts to that side and produces thrust in that direction, causing the

helicopter to hover sideways.



2. Collective. The collective pitch control, or collective, is located on the left side of the

pilot’s seat with a pilot-selected variable friction control to prevent inadvertent movement. The

collective changes the pitch angle of all the main rotor blades collectively (i.e., all at the same

time) and independently of their positions. Therefore, if a collective input is made, all the blades

change equally, increasing or decreasing total lift or thrust, with the result of the helicopter

increasing or decreasing in altitude or airspeed.

3. Anti-torque Pedals. The anti-torque pedals are located in the same position as the rudder

pedals in a fixed-wing aircraft and serve a similar purpose, namely to control the direction in

which the nose of the aircraft is pointed. Application of the pedal in a given direction changes

the pitch of the tail rotor blades, increasing or reducing the thrust produced by the tail rotor,

causing the nose to yaw in the direction of the applied pedal. The pedals mechanically change

the pitch of the tail rotor, altering the amount of thrust produced.

4. Throttle. Helicopter rotors operate at a specific RPM. The throttle controls the power

produced by the engine, which connects to the rotor by a transmission. The purpose of the

throttle is to maintain enough engine power to keep the rotor RPM within allowable limits to

produce enough lift for flight. In single-engine helicopters, if so equipped, the throttle control is

typically a twist grip mounted on the collective control, but it can also be a lever mechanism in

fully governed systems, as seen in the TH-57. Multi-engine helicopters generally have a power

lever or mode switch for each engine.

105. FLIGHT CONDITIONS

There are two basic flight conditions for a helicopter: hover and forward flight. Hovering is the

most challenging part of flying a helicopter. This is because a helicopter generates its own gusty

air while in a hover, which acts against the fuselage and flight control surfaces. The result is the

need for constant control inputs and corrections by the pilot to keep the helicopter where it is

required to be. Despite the complexity of the task, the control inputs in a hover are simple. The

cyclic is used to eliminate drift in the horizontal direction: forward, backward, right and left.

The collective is used to maintain altitude. The pedals are used to control nose direction or

heading. The interaction of these controls makes hovering so difficult, since an adjustment in

any one control requires an adjustment of the other two, creating a cycle of constant correction.

Displacing the cyclic forward initially causes the nose to pitch down, with a resultant increase in

airspeed and loss of altitude. Aft cyclic initially causes the nose to pitch up, slowing the

helicopter and causing it to climb; however, as the helicopter reaches a state of equilibrium, the

horizontal stabilizer helps level the helicopter to minimize drag, unlike an airplane. Therefore,

the helicopter has very little pitch deflection up or down when the helicopter is stable in a flight

mode. The variation from absolutely level attitude depends on the particular helicopter and the

horizontal stabilizer function.

Increasing collective (power) while maintaining a constant airspeed induces a climb while

decreasing collective causes a descent. Coordinating these two inputs, down collective plus aft

cyclic or up collective plus forward cyclic, results in airspeed changes while maintaining a

constant altitude.



The pedals serve the same function in both a helicopter and a fixed-wing aircraft, to maintain

balanced flight. This is done by applying pedal input in whichever direction is necessary to

center the ball in the turn and bank indicator.

THE ATMOSPHERE 2-1

CHAPTER TWO

THE ATMOSPHERE

200. INTRODUCTION

The purpose of this chapter is to provide a review of basic physics and atmospherics relevant to

helicopter aerodynamics. Review Fundamentals of Aerodynamics Student Guide

(NAVAVSCOLSCOM-SG-111) from API if necessary.

201. LESSON TOPIC LEARNING OBJECTIVES







202. REVIEW OF BASIC PHYSICS AND AERODYNAMICS

Based on the student naval aviator’s recent completion of aviation preflight indoctrination,

vector analysis, physics and atmospherics is abbreviated and includes only those concepts

applicable to rotary wing flight. If necessary, review “Fundamentals of Aerodynamics”

workbook from API. A few concepts, such as density altitude (DA) computation, are introduced

in this section as it relates to the reviewed material.

203. VECTOR ANALYSIS

A scalar is a quantity that describes only magnitude, e.g., time, temperature, or volume. It is

expressed as a single number including units. A vector is a quantity that describes both

magnitude and direction. It commonly represents displacement, velocity, acceleration, or force.

Vectors are represented as arrows. The direction and length of the arrow represent the direction

and magnitude of the vector. Vectors may be added by placing the tail of each succeeding vector

on the head (or tip) of the one preceding it and drawing the resultant vector from the tail of the

first to the tip of the last. This new vector is the resulting magnitude and direction of all the

original vectors working together.

Conversely, a vector may be deconstructed into two or more component vectors that lie in a

desired plane of motion or direction.

CHAPTER TWO HELICOPTER AERODYNAMICS WORKBOOK

2-2 THE ATMOSPHERE

Figure 2-1 Resultant by the Tip-to-Tail Method

Figure 2-2 Force Vectors on an Airfoil

Figure 2-3 Force Vectors on Aircraft in Flight

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER TWO

THE ATMOSPHERE 2-3

Figures 2-2 and 2-3 show an example of vectors used to depict the forces acting on an airfoil

segment, and an aircraft in flight. Note that the total rotor thrust vector may be resolved into

perpendicular components, a vertical component and a horizontal component. The drag and

weight can also combine to form a resultant of these two forces, which must then be equal and

opposite to the total rotor thrust in equilibrium flight.

204. ESSENTIAL TERMS AND DEFINITIONS

1. Volume (v) is the amount of space occupied by an object.

2. Density ( or ‘rho’) is mass per unit volume.

3. Weight (W) is the force with which a mass is attracted toward the center of the earth by

gravity.

4. A moment (M) is created when a force is applied at some distance from an axis, and tends

to produce rotation about that point. A moment is a vector quantity equal to a force (F) times

the distance (d) from the point of rotation on a line that is perpendicular to the applied

force vector. This perpendicular distance is called the moment arm. Torque (Q) is another

word for a moment created by a force.

5. Work (W) is done when a force acts on a body and moves it. It is a scalar quantity equal

to the force (F) times the distance of displacement (s).

6. Power (P) is the rate of doing work or work done per unit of time.

7. Energy is a scalar measure of a body’s capacity to do work. There are two types of

energy: potential energy and kinetic energy. Energy cannot be created or destroyed, but may be

transformed from one form to another. This principle is called the law of conservation of energy.

8. Potential energy (P.E.) is the ability of a body to do work because of its position or state

of being. It is a function of mass (m), gravity (g), and height (h).

9. Kinetic energy (K.E.) is the ability of a body to do work because of its motion. It is a

function of mass (m) and velocity (V).

Work may be performed on a body to change its position and give it potential energy or work

may give the body motion so that it has kinetic energy. Under ideal conditions, if no work is

being done on an object, its total energy will remain constant. Such an object is considered to be

in a closed system. In a closed system, the total energy will remain constant but potential energy

may be converted to kinetic energy, and vice versa. For example, the kinetic energy of a glider

in forward flight may be converted into potential energy by climbing. As the glider’s altitude

(P.E.) increases, its velocity (K.E.) will decrease.


2-4 THE ATMOSPHERE

205. PROPERTIES OF THE ATMOSPHERE

The atmosphere is composed of approximately 78% nitrogen (N2), 21% oxygen (O2), and 1%

other gases by volume, which includes argon and carbon dioxide. Air is considered a uniform

mixture of these gases and will be examined as a whole rather than as separate gases.

1. Static pressure (PS) is the force each air particle exerts on those around it. On a more

macroscopic scale, ambient static pressure (14.7 psi at sea level on standard day) is equal to the

weight of a column of air over a given area. The force of static pressure acts perpendicularly to

any surface with which the air particles collide. As altitude increases, less air is above you, so

the weight of the column of air decreases. Thus atmospheric static pressure decreases with an

increase in altitude at a rate of approximately 1.0 in-Hg per 1000 feet, near the earth’s surface.

Figure 2-4 Static pressure

2. Air density ( ) is the total mass of air particles per unit of volume. The distance between

individual air particles increases with altitude resulting in fewer particles per unit volume.

Therefore, air density decreases with an increase in altitude.

3. Density Ratio (σ) is the ratio of the density of air at a specific altitude to that of the

standard altitude (sea level).

4. Temperature (T) is a measure of the average kinetic energy of air particles. As

temperature increases, particles begin to move and vibrate faster, increasing their kinetic energy.


THE ATMOSPHERE 2-5

Air temperature decreases linearly with an increase in altitude at a rate of approximately

2 °C (3.57 F) per 1000 ft. up through 36,000 feet MSL. This is called the standard, or

adiabatic lapse rate. Above 36,000 feet lies the isothermal layer where air is at a constant

temperature of -56.5 °C.

Figure 2-5 Molecular Energy and Air Temperature

5. Humidity is the amount of water vapor in the air. As humidity increases, water molecules

displace an equal number of air molecules. Since water molecules have less mass (H2O,

molecular weight (MW) 18) than air (N2, MW 28; and O2, MW 32) and occupy approximately

the same volume, the overall mass in a given volume decreases. Therefore, as humidity

increases, air density decreases. Compared to dry air, the density of air at 100% humidity is

4% less.

6. Viscosity () is a measure of the air's resistance to flow and shearing. Air viscosity can

determine its tendency to either stick to a surface or how easily it flows past it. For liquids, as

temperature increases, viscosity decreases. Recall that the oil in your car flows better or “gets

thinner” when the engine gets hot. Just the opposite happens with air: Air viscosity increases

with an increase in temperature.

206. THE STANDARD ATMOSPHERE

The atmospheric layer were most flying is done is an ever-changing environment. Temperature

and pressure vary with altitude, season, location, time, and even solar sunspot activity. It is

impractical to take all of these into consideration when discussing aircraft performance. In order

to disregard these atmospheric changes, an engineering baseline was developed called the

standard atmosphere. It is a set of reference conditions giving average values of air properties

as a function of altitude. Unless otherwise stated, any discussion of atmospheric properties

in this course will assume standard atmospheric conditions.


2-6 THE ATMOSPHERE

Figure 2-6 Standard Atmospheric Table

207. ALTITUDE COMPUTATIONS

One of the benefits of a standard atmosphere is the concepts of pressure altitude (PA) and

density altitude (DA). PA is that altitude in the standard atmosphere that corresponds to a

particular static air pressure. An aircraft altimeter senses pressure through the static portion of

the pitot-static system, then shows the altitude at which that pressure would be found in the

standard atmosphere. Early altimeters were referenced to the standard Sea Level pressure of

29.92 inches Hg at sea level. Altitude was estimated by determining the altitude in the standard

atmosphere at which the measured pressure would occur. Modern altimeters are adjusted to

yield accurate altitude at a known point, like an airfield, by changing reference sea level pressure

to an appropriate value. Airport meteorologists measure air pressure, determine what pressure at

sea level would have to exist to yield an accurate altitude reading using standard pressure lapse

rate, and report that setting to pilots. After being set properly, an adjustable altimeter reports the

altitude at which the measured pressure would be found if a standard pressure lapse rate was

applied to a sea level pressure equal to that set by the pilot. Standard pressure lapse rate is

1000 feet of PA for each one inch of Hg. To determine PA (in the standard atmosphere) for use

in calculations a pilot needs only to set 29.92 inches into the altimeter and read the result.


THE ATMOSPHERE 2-7

Properly setting the altimeter and periodically checking for changes in the prescribed setting can

be very important. To properly set the altimeter, pilots should make sure that they are receiving

the altimeter setting in inches of mercury for height above sea level. Three types of settings may

be encountered overseas: QNH, QNE, and QFE.

1. QNH is altitude corrected to standard sea level (provided in ATIS in the United States).

2. QNE is PA (altimeter set to 29.92)

3. QFE is a setting at an airfield to read height above ground at that location (the altimeter

would read zero at the airfield surface).

Overseas, settings are often given in hectopascals (millibars), rather than inches of mercury. To

avoid problems, listen for foreign controllers’ statements of what their pressure reference is and,

if necessary, use the Flight Information Handbook to convert.

Checking for changes in altimeter setting during cross country travel or as weather moves in can

also be important. When station pressure or temperature drops, an altimeter set at the previous

condition will read higher than it should. For example, a flight at 1000 ft. AGL in the standard

atmosphere would have the aircraft at a pressure of 28.92. Note that flight into a high pressure

system would cause the altimeter to read low (actual altitude of 2000 ft. AGL) while flight into a

low pressure system would cause the altimeter to read higher than actual altitude (0 ft. AGL).

REMEMBER: “HIGH TO LOW LOOK OUT BELOW” OR “HOT TO COLD LOOK

OUT BELOW.”

Figure 2-7 Pressure Altitude vs. True Altitude with Varying Local Pressures


2-8 THE ATMOSPHERE

208. ALTITUDE COMPUTATIONS – DENSITY ALTITUDE

A more appropriate term for correlating aerodynamic performance in the nonstandard

atmosphere is density altitude (DA). DA is that altitude in the standard atmosphere which

corresponds to a particular air density. Changes in air density are caused by variations in

atmospheric pressure, temperature, and humidity. Density altitude (DA) is the PA corrected for

temperature and humidity deviations from the standard atmosphere. If you know the

outside air temperature and humidity, and you can determine the PA, you can derive the DA.

DA is the environmental factor that most significantly affects power available.

Typically, aviators use a chart to determine DA for the ambient PA and temperature. Using

Figure 2-8, enter the chart at the bottom at the appropriate outside air temperature (OAT) and

plot vertically upward to intersect the current PA depicted on the diagonals, determined by

dialing 29.92 into the aircraft altimeter. From this point, read laterally to the left to determine the

DA (not corrected for humidity). In the example depicted, a temperature of 6 °C and 2400’ PA,

results in a 2000’ DA.

Another method for estimating DA is to use the ‘rule of thumb’ equation below.

DA = PA + [(TAmbient – TStd@Altitude)] x 120

For PA, dial in 29.92 " Hg. Temperatures are in C.

The above ‘rule of thumb’ merely requires an accurate understanding of standard temperature at

altitude. Remember that the temperature at sea level for a standard day is 15 °C. With an

average lapse rate of 2 °C / 1000’ MSL, the standard temperature at altitude can be easily

determined; i.e., 5 °C at 5000’ MSL, and -5 °C at 10,000’ MSL. This estimation of DA still fails

to correct for humidity, which is discussed below.

As mentioned earlier, changes in the water vapor content, or humidity, can also greatly affect

the density of the air, in addition to temperature deviations from standard. To recap, as humidity

increases, water molecules with less mass and approximately the same volume as air molecules

displace the more dense air molecules to make the same overall volume containing less actual

mass. Thus, an increase in humidity leads to a decrease in air density, and, therefore, an increase

in DA.


THE ATMOSPHERE 2-9

Figure 2-8 DA Chart

One way to adjust calculations for humidity is to use a higher temperature than might be

associated with lower density in performance charts. This fictional quantity is known as virtual

temperature, and is defined as OAT corrected for relative humidity. In the same way that wind

chill is applied to a cold day’s temperature to reflect how the wind affects the human body, a

virtual temperature correction may be applied to a temperature measurement to reflect the effect

of humidity on the air’s density. The dew point temperature correction chart accepts dew point

and temperature, and then yields virtual temperature and DA.


2-10 THE ATMOSPHERE

Figure 2-9 Dew Point Correction Chart

Moisture in the air can be slightly beneficial in controlling engine temperature, but generally

tends to be detrimental to helicopter performance. The decrease in the density of the airflow due

to the presence of water molecules requires more mass flow of air to produce the same amount of

thrust. Since only a limited amount of air can be run through the engine, the effect of water

vapor is to reduce power available.

This loss of power available and aforementioned cooling effects tend to offset each other in

typical helicopter operating scenarios, so engine performance doesn’t change as much with

humidity as it does with temperature.

With respect to rotor systems, reduced air density decreases the lift produced on the rotor blades.

For this reason, the overall effect of humidity degrades helicopter performance.

A common adjustment is the 10% rule of thumb. Add 100 feet to your DA (based on PA and

OAT), for every 10% relative humidity above 0% RH.

DA (corrected for RH) = DA (chart) + (100’ x RH/10%)

The 10% adjustment factor stems from a linear approximation of the curve. Some interpretations

of the curve state that the relative humidity (RH) correction (10% adjustment factor) doesn't go


THE ATMOSPHERE 2-11

into effect until the RH is above 40% (40% rule). This is because the same curve in Figure 2-10

can be estimated by two distinct slopes, with this 40% rule appearing to give a better

approximation. For example, 0-40% humidity results in no correction, and 50-100% humidity

results in a 100-600 foot altitude correction for humidity (rather than an errant 500-1000 foot

correction) based on the “10% rule.”

Figure 2-10 Thrust Variation with Humidity

A comparison of the three commonly used methods to compute DA indicates that a significant

variation exists between the chosen methods. For example, if we assume an OAT of 30 C

(86 F), PA of sea level and Dew Point of 30 C (RH=100%) we obtain a DA of 1800 feet based

on the DA chart using PA and an uncorrected OAT. If we incorporate the "100 foot-10% RH

rule" for 100% RH the DA is increased by 1000 feet to 2800 feet or to 2400 feet if we only apply

the "rule of thumb" for RH above 40% as some NATOPS manuals dictate. The actual DA

obtained from the DA chart using a corrected OAT yields 2200 feet.

Note the effect of the differences applied to typical aircraft performance data (CH53D) in

Figure 2-11. Making no adjustment might lead to overestimation of capabilities. The difference

is relatively small, but may result in operations conducted close to safe power margins. Ignoring

DA effects could result in exceeding safe margins. Being overly conservative could limit the

ability to complete a mission satisfactorily. The 40% rule offers the most accurate quick

estimate of humidity effects, compared to the more conservative 10% rule.


2-12 THE ATMOSPHERE

DA (corrected for RH >40%) = DA (chart) + 100’ (RH – 40%)/10%

And is only corrected for RH > 40%

Again, for calculations in the field, PA is easily obtained by setting 29.92" Hg in the Kollsman

window of the barometric altimeter. Obtaining the dew point is usually a little more difficult

unless you have access to a weather service. In the absence of any of this information, you can

always assume a worst case scenario of 100% RH which is when the OAT and dew point are the

same.

For comparison purposes, Figure 2-11 depicts for a CH-53D the relative accuracy of the three

techniques for humidity correction of DA with an RH of 100%.

Adjustment DA Max Gross Wt. (HOGE)

NONE 7800 FT 34,800 lbs.

10% RULE 8800 FT 34,000 lbs.

DA COMPUTER 8400 FT 34,500 lbs.

40% RULE 8400 FT 34,500 lbs.

Figure 2-11 Sample Effects of DA Calculations on CH-53D Performance

The 10% Rule provides a more conservative estimate of DA and is the recommended method in

most helicopter NATOPS manuals. If a NATOPS manual does not discuss the effects of

humidity, be conservative and apply the 10% Rule.

HELICOPTER AERODYNAMICS BASICS 3-1

CHAPTER THREE

HELICOPTER AERODYNAMICS BASICS

300. INTRODUCTION

The purpose of this chapter is to introduce the student to aerodynamic principles and

fundamentals that apply to helicopters.












302. AIRCRAFT REFERENCE SYSTEM

An aircraft's (helicopter or fixed-wing) reference system consists of three mutually perpendicular

lines (axes) intersecting at a single point. This point, called the center of gravity (CG), is the

point at which all weight is concentrated and at which all forces are measured. Theoretically, the

aircraft will balance if suspended at the CG. When in flight, the aircraft will rotate about the CG,

so all moments will be resolved around it as well. The CG will move as fuel burns,

bombs/missiles expended, or cargo shifts.

The longitudinal axis passes from the nose to the tail of the aircraft. Rotation on the

longitudinal axis is roll, or lateral control. Movement along the longitudinal axis is surge.

The lateral axis passes from wingtip to wingtip. Rotation on the lateral axis is pitch, or

longitudinal control. Movement along the lateral axis is sway.

CHAPTER THREE HELICOPTER AERODYNAMICS WORKBOOK

3-2 HELICOPTER AERODYNAMICS BASICS

The vertical axis passes vertically through the CG. Rotation on the vertical axis is yaw, or

directional control. Movement along the vertical axis is heave. As an aircraft moves through

the air, the axis system moves with it. Therefore, the movement of the aircraft can be described

by the movement of its CG.

Figure 3-1 Aircraft Reference System

303. GENERAL

Gravity acting on the mass (the amount of matter) of an object creates a force called weight. The

rotor blade in Figure 3-2 weighs 100 lbs. It is 20 feet long (span) and is 1 foot wide (chord).

Accordingly, its surface area is 20 square feet.

The blade is perfectly balanced on a pinpoint stand, as seen in Figure 3-3, from looking at it from

the end (the airfoil view). The goal is for the blade to defy gravity and stay exactly where it is

when we remove the stand. If we do nothing before removing the stand, the blade will simply

fall to the ground. Can we exert a force (a push or pull) opposite gravity that equals the 100 lb.

weight of the blade? Yes, in helicopters we use aerodynamic force to oppose weight and to

maneuver.

Every object in the atmosphere is surrounded by a gas that exerts a static force of 2,116 lb. per

square foot (a force times a unit area, called pressure) at sea level. However, that pressure is

exerted equally all over the blade (top and bottom) and therefore does not create any useful force

on the blade. We need only create a difference of a single pound of static pressure differential

per square foot of blade surface to have a force equal to the blade’s weight (100 lb. of upward

pressure opposite 100 lb. downward weight).

Total pressure consists of static pressure and, if the air is moving, dynamic pressure (a pressure

in the direction of the air movement). If dynamic pressure is increased the static pressure will

decrease. Due to the design of the airfoil, the velocity of the air passing over the upper surface

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER THREE


will be greater than that of the lower surface, leading to higher dynamic pressure on the upper

surface than on the lower surface. The higher dynamic pressure on the upper surface lowers the

static pressure on the upper surface. The static pressure on the bottom will now be greater than

the static pressure on the top. The blade will experience an upward force. With just the right

amount of air passing over the blade the upward force will equal one pound per square foot.

This upward force is equal to, and acts opposite the blade’s weight of 100 lb. So, if we now

remove the stand, the blade will defy gravity and remain in its position (ignoring rearward drag

for the moment).

Figure 3-2 Area of a Blade

Figure 3-3 Profile of an airfoil

The force created by air moving over an object (or moving an object through the air) is called

aerodynamic force. Aero means air. Dynamic means moving or motion. Accordingly, by

moving the air over an airfoil we can change the static pressures on the top and bottom thereby

generating a useful force (an aerodynamic force). The portion of the aerodynamic force that is

usually measured perpendicular to the air flowing around the airfoil is called lift and is used to

oppose weight. Drag is the portion of aerodynamic force that is measured as the resistance

created by an object passing through the air (or having the air passed over it). Drag acts in a

stream-wise direction with the wind passing over the airfoil and retards forward movement.



Figure 3-4 Forces acting on a helicopter in forward flight

304. FORCES ACTING ON THE AIRCRAFT

Once a helicopter leaves the ground, it is acted upon by four aerodynamic forces; thrust, drag,

lift, and weight. Understanding how these forces work and knowing how to control them with

the use of power and flight controls are essential to flight. They are defined as follows:

Lift: opposes the downward force of weight, is produced by the dynamic effect of the air acting

on the airfoil and acts perpendicular to the flightpath through the center of lift.

Weight: the combined load of the aircraft itself, the crew, the fuel, and the cargo or baggage.

Weight pulls the aircraft downward because of the force of gravity. It opposes lift and acts

vertically downward through the aircraft’s center of gravity (CG).

Thrust: the force produced by the power plant/ propeller or rotor. It opposes or overcomes the

force of drag. As a general rule, it acts parallel to the longitudinal axis. However, this is not

always the case, as explained later.

Drag: a rearward, retarding force caused by disruption of airflow by the wing, rotor, fuselage,

and other protruding objects. Drag opposes thrust and acts rearward parallel to the relative wind.

305. LIFT

Lift is generated when an object changes the direction of flow of a fluid or when the fluid is

forced to move by the object passing through it. When the object and fluid move relative to each



other and the object turns the fluid flow in a direction perpendicular to that flow, the force

required to do this work creates an equal and opposite force that is lift. The object may be

moving through a stationary fluid, or the fluid may be flowing past a stationary object. These

two are effectively identical as, in principle, it is only the frame of reference of the viewer which

differs. The lift generated by an airfoil depends on such factors as:

1. Speed of the airflow

2. Density of the air

3. Total area of the segment or airfoil

4. Angle of attack (AOA) between the air and the airfoil

The AOA is the angle at which the airfoil meets the oncoming airflow (or vice versa). In the

case of a helicopter, the object is the rotor blade (airfoil) and the fluid is the air. Lift is produced

when a mass of air is deflected, and it always acts perpendicular to the resultant relative wind. A

symmetric airfoil must have a positive AOA to generate positive lift. At a zero AOA, no lift is

generated. At a negative AOA, negative lift is generated. A cambered or nonsymmetrical airfoil

may produce positive lift at zero, or even small negative AOA.

The basic concept of lift is simple. However, the details of how the relative movement of air and

airfoil interact to produce the turning action that generates lift are complex. In any case causing

lift, an angled flat plate, revolving cylinder, airfoil, etc., the flow meeting the leading edge of the

object is forced to split over and under the object. The sudden change in direction over the

object causes an area of low pressure to form behind the leading edge on the upper surface of the

object. In turn, due to this pressure gradient and the viscosity of the fluid, the flow over the

object is accelerated down along the upper surface of the object. At the same time, the flow

forced under the object is rapidly slowed or stagnated causing an area of high pressure. This also

causes the flow to accelerate along the upper surface of the object. The two sections of the fluid

each leave the trailing edge of the object with a downward component of momentum, producing

lift.

Figure 3-5 Production of lift



Bernoulli’s Principle

Bernoulli’s principle describes the relationship between internal fluid pressure and fluid velocity.

It is a statement of the law of conservation of energy and helps explain why an airfoil develops

an aerodynamic force. The concept of conservation of energy states energy cannot be created or

destroyed and the amount of energy entering a system must also exit. Specifically, in this case

the “energy” referred to is the dynamic pressure (the kinetic energy of the air - more velocity,

more kinetic energy) and static air pressure (potential energy). These will change among

themselves, but the total pressure energy remains constant inside the tube.

A simple tube with a constricted portion near the center of its length illustrates this principle. An

example is running water through a garden hose. The mass of flow per unit area (cross-sectional

area of tube) is the mass flow rate. In Figure 3-6 the flow into the tube is constant, neither

accelerating nor decelerating; thus, the mass flow rate through the tube must be the same at

stations 1, 2, and 3. If the cross-sectional area at any one of these stations, or any given point, in

the tube is reduced, the fluid velocity must increase to maintain a constant mass flow rate to

move the same amount of fluid through a smaller area. The continuity of mass flow causes the

air to move faster through the venturi. In other words, fluid speeds up in direct proportion to the

reduction in area.

Figure 3-6 Water flow through a tube

Bernoulli (Ptotal = Pdynamic + Pstatic) states that the increase in velocity will increase the stream-wise

dynamic pressure. Since the total pressure in the tube must remain constant, the static pressure

on the sides of the venturi will decrease. Venturi effect is the term used to describe this

phenomenon.

Figure 3-7 illustrates plates of one square foot in the dynamic flow and on the sides of the tube

indicating static pressure, with corresponding pressure. At point 2, it is easier to visualize the

static pressure reduction on the top of the airfoil as compared to the bottom of the airfoil, which

is depicted as outside of the tube and therefore at ambient static pressure. Keep in mind with

actual blades it is not a simple as this example because the bottom static pressure is influenced

by blade design and blade angle, among other things. However, the basic idea is that it is the

static pressure differential between the top and bottom multiplied by the surface area of the blade

that generates the aerodynamic force.



Figure 3-7 Venturi effect

Venturi Flow

While the amount of total energy within a closed system (the tube) does not change, the form of

the energy may be altered. Pressure of flowing air may be compared to energy in that the total

pressure of flowing air always remains constant unless energy is added or removed. Fluid flow

pressure has two components, static and dynamic pressure. Static pressure is the pressure

component measured in the flow but not moving with the flow as pressure is measured. Static

pressure is also known as the force per unit area acting on a surface. Dynamic pressure of flow

is that component existing as a result of movement of the air. The sum of these two pressures is

total pressure. As air flows through the constriction, static pressure decreases as velocity

increases. This increases dynamic pressure. Figure 3-7 depicts the bottom half of the constricted

area of the tube, which resembles the top half of an airfoil. Even with the top half of the tube

removed, the air still accelerates over the curved area because the upper air layers restrict the

flow, just as the top half of the constricted tube did. This acceleration causes decreased static

pressure above the curved portion and creates a pressure differential caused by the variation of

static and dynamic pressures.

Newton’s Third Law of Motion

Additional lift is provided by the rotor blade’s lower surface as air striking the underside is

deflected downward. According to Newton’s Third Law of Motion, “for every action there is an

equal and opposite reaction,” the air that is deflected downward also produces an upward

(lifting) reaction.

Since air is much like water, the explanation for this source of lift may be compared to the

planing effect of skis on water. The lift that supports the water skis (and the skier) is the force

caused by the impact pressure and the deflection of water from the lower surfaces of the skis.

Under most flying conditions, the impact pressure and the deflection of air from the lower

surface of the rotor blade provides a comparatively small percentage of the total lift. The

majority of lift is the result of decreased pressure above the blade, rather than the increased

pressure below it.



306. WEIGHT

Normally, weight is thought of as being a known, fixed value, such as the weight of the

helicopter, fuel, and occupants. To lift the helicopter off the ground vertically, the rotor disk

must generate enough lift to overcome or offset the total weight of the helicopter and its

occupants. Newton’s First Law states: “Every object in a state of uniform motion tends to

remain in that state of motion unless an external force is applied to it.” In this case, the object is

the helicopter whether at a hover or on the ground and the external force applied to it is lift,

which is accomplished by increasing the pitch angle of the main rotor blades. This action forces

the helicopter into a state of motion, without it the helicopter would either remain on the ground

or at a hover.

The weight of the helicopter can also be influenced by aerodynamic loads. When you bank a

helicopter while maintaining a constant altitude, the “G” load or load factor increases. The load

factor is the actual load on the rotor blades at any time, divided by the normal load or gross

weight (weight of the helicopter and its contents). Any time a helicopter flies in a constant

altitude curved flightpath, the load supported by the rotor blades is greater than the total weight

of the helicopter. The tighter the curved flightpath is, the steeper the bank is; the more rapid the

flare or pullout from a dive is, the greater the load supported by the rotor. Therefore, the greater

the load factor must be.

To overcome this additional load factor, the helicopter must be able to produce more lift. If

excess engine power is not available, the helicopter either descends or has to decelerate in order

to maintain the same altitude. The load factor and, hence, apparent gross weight increase is

relatively small in banks up to 30°. Even so, under the right set of adverse circumstances, such

as high-density altitude, turbulent air, high gross weight, and poor pilot technique, sufficient or

excess power may not be available to maintain altitude and airspeed. Pilots must take all of these

factors into consideration throughout the entire flight from the point of ascending to a hover to

landing. Above 30° of bank, the apparent increase in gross weight soars. At 30° of bank, or

pitch, the apparent increase is only 16 percent, but at 60°, it is twice the load on the wings and

rotor disk. For example, if the weight of the helicopter is 1,600 pounds, the weight supported by

the rotor disk in a 30° bank at a constant altitude would be 1,856 pounds (1,600 + 16 percent (or

256)). In a 60° bank, it would be 3,200 pounds; in an 80° bank, it would be almost six times as

much, or 8,000 pounds. It is important to note that each rotor blade must support a percentage of

the gross weight. In a two-bladed system, each blade of the 1,600-pound helicopter as stated

above would have to lift 50 percent or 800 pounds. If this same helicopter had three rotor blades,

each blade would have to lift only 33 percent, or 533 pounds. One additional cause of large load

factors is rough or turbulent air. The severe vertical gusts produced by turbulence can cause a

sudden increase in AOA, resulting in increased rotor blade loads that are resisted by the inertia of

the helicopter.

Each type of helicopter has its own limitations that are based on the aircraft structure, size, and

capabilities. Regardless of how much weight one can carry or the engine power that it may have,

they are all susceptible to aerodynamic overloading. Unfortunately, if the pilot attempts to push

the performance envelope the consequence can be fatal. Aerodynamic forces effect every

movement in a helicopter, whether it is increasing the collective or a steep bank angle.



Anticipating results from a particular maneuver or adjustment of a flight control is not good

piloting technique. Instead pilots need to truly understand the capabilities of the helicopter under

any and all circumstances and plan never to exceed the flight envelope for any situation.

Figure 3-8 Notional load factor diagram

307. THRUST

Thrust, like lift, is generated by the rotation of the main rotor disk. In a helicopter, thrust can be

forward, rearward, sideward, or vertical. The resultant lift and thrust determines the direction of

movement of the helicopter.

The solidity ratio is the ratio of the total rotor blade area, which is the combined area of all the

main rotor blades, to the total rotor disk area. This ratio provides a means to measure the

potential for a rotor disk to provide thrust and lift. The mathematical calculations needed to

calculate the solidity ratio for each helicopter may not be of importance to most pilots but what

should be are the capabilities of the rotor disk to produce and maintain lift. Many helicopter

accidents are caused from the rotor disk being overloaded. Simply put, pilots attempt maneuvers

that require more lift than the rotor disk can produce or more power than the helicopter’s power-

plant can provide. Trying to land with a nose high attitude along with any other unfavorable

condition (i.e., high gross weight or wind gusts) is most likely to end in disaster.

The tail rotor also produces thrust. The amount of thrust is variable through the use of the anti-

torque pedals and is used to control the helicopter’s yaw.

308. DRAG

The force that resists the movement of a helicopter through the air and is produced when lift is

developed is called drag. Drag must be overcome by the engine to turn the rotor. Drag always

acts parallel to the relative wind. Total drag is composed of three types of drag: profile, induced,

and parasite.



Figure 3-9 Notional Drag profiles

Profile Drag

Profile drag develops from the frictional resistance of the blades passing through the air. It does

not change significantly with the airfoil’s AOA but increases moderately when airspeed

increases. Profile drag is composed of form drag and skin friction. Form drag results from the

turbulent wake caused by the separation of airflow from the surface of a structure. The amount

of drag is related to both the size and shape of the structure that protrudes into the relative wind.

Skin friction is caused by surface roughness. Even though the surface appears smooth, it may be

quite rough when viewed under a microscope. A thin layer of air clings to the rough surface and

creates small eddies that contribute to drag.

Induced Drag

Induced drag is generated by the airflow circulation around the rotor blade as it creates lift. The

high-pressure area beneath the blade joins the low-pressure area above the blade at the trailing

edge and at the rotor tips. This causes a spiral, or vortex, which trails behind each blade

whenever lift is being produced. These vortices deflect the airstream downward in the vicinity of

the blade, creating an increase in downwash. Therefore, the blade operates in an average relative



wind that is inclined downward and rearward near the blade. Because the lift produced by the

blade is perpendicular to the relative wind, the lift is inclined aft by the same amount. The

component of lift that is acting in a rearward direction is induced drag.

Figure 3-10 Induced Drag

As the air pressure differential increases with an increase in AOA, stronger vortices form, and

induced drag increases. Since the blade’s AOA is usually lower at higher airspeeds, and higher

at low speeds, induced drag decreases as airspeed increases and increases as airspeed decreases.

Induced drag is the major cause of drag at lower airspeeds.

Parasite Drag

Parasite drag is present any time the helicopter is moving through the air. This type of drag

increases with airspeed. Non-lifting components of the helicopter, such as the cabin, rotor mast,

tail, and landing gear, contribute to parasite drag. Any loss of momentum by the airstream, due

to such things as openings for engine cooling, creates additional parasite drag. Because of its

rapid increase with increasing airspeed, parasite drag is the major cause of drag at higher

airspeeds. Parasite drag varies with the square of the velocity; therefore, doubling the airspeed

increases the parasite drag four times.

Total Drag

Total drag for a helicopter is the sum of all three drag forces. As airspeed increases, parasite

drag increases, while induced drag decreases. Profile drag remains relatively constant

throughout the speed range with some increase at higher airspeeds. Combining all drag forces

results in a total drag curve. The low point on the total drag curve shows the airspeed at which

drag is minimized. This is the point where the lift-to-drag ratio is greatest and is referred to as

L/DMAX. At this speed, the total lift capacity of the helicopter, when compared to the total drag

of the helicopter, is most favorable. This is an important factor in helicopter performance.



Figure 3-11 Drag Curve

309. INTRODUCTION TO LIFT THEORIES

Several theories developed during the last two hundred years to attempt to explain the production

of lift by an airfoil: Pressure Distribution, Circulation, Momentum, and the Blade Element

Theory. One argument says that the pressure differential from the top to bottom of the airfoil

describes lift. Another argument says that the wing deflects the air downward, thus pushing

itself up. Yet another says that there is a net circulation of air around the wing, which causes it

to lift. Each of these ideas is completely legitimate, supported by mathematical proof. Each also

has a set of specific constraints and approximations, which limit its applicability. There will be

situations where one method is more convenient to use than another and other situations that

demand one specific method of investigation.

310. PRESSURE DISTRIBUTION THEORY

The pressure distribution theory evolves from the principle of continuity, and the principle of

conservation of energy as applied to fluid dynamics (Bernoulli’s Equation). To recap, in

considering continuity, as the area (of a stream-tube, for example) decreases, the velocity

increases. Furthermore, from Bernoulli, as the velocity of the air goes up, the static pressure

goes down.



Therefore, putting the Bernoulli equation together with the continuity principle, we have

the following: as the area decreases, the velocity increases, and as the velocity increases the

static pressure decreases. If the pressure goes down over the top of a wing more so than over

the bottom, then the wing will be lifted up as seen in Figure 3-12. This differential pressure,

accounted for by the continuity principle and the Bernoulli equation, is the method of choice in

describing the mechanics of lift by the pressure distribution theory.

Figure 3-12 Pressure Changes Around a Cambered Airfoil

311. CIRCULATION THEORY

Circulation theory, or the Kutta-Joukowski Theorem, is a method for describing the flow over a

spinning cylinder and, more generally, over any closed area (Figure 3-13). If a non-rotating

circular cylinder is placed in a flow field it will produce no lift. The streamlines and resultant

pressure distributions around a cylinder without circulation (Figure 3-14) generate no net lift

force. When the cylinder is rotated, however, it induces a rotational or circulatory flow and there

is a distinct change in the streamlines and pressure distributions. Air next to the surface of the

cylinder is sped up on the top and slowed down on the bottom by the relative motion of the

cylinder’s surface. The differences in flow speeds cause pressure differences on top and bottom

(Figure 3-14), with the end result being a net lift force perpendicular to the relative velocity.

There have been recent developments in rotary wing flight where circulation theory has shown

nearly direct application. Use of a rotating body to generate pressure differences on top and

bottom surfaces, or viewed alternatively, impart circulation to a flow, is termed Magnus effect.



Figure 3-13 Magnus Effect

Figure 3-14 Pressure Distribution for Magnus Effect

312. MOMENTUM THEORY

The Momentum Theory of lift relies on Newton’s laws of motion with regard to the air as it

passes over an airfoil and the reaction of the airfoil to the motion of the air. In a hover, this

theory states that a certain amount of air above the rotor system is accelerated to a certain

velocity at a certain distance below the rotor. Since the amount of air has a finite mass and is

given a finite acceleration, its force can be determined through Newton’s second law (F=ma).

Specifically, the theory shows that given an initial velocity (v0) of zero well above the rotor

system, the rotor system accelerates the air downward through the rotors to a particular velocity

(vi, induced velocity) based on the diameter of the rotor, the density of the air and the weight of



the helicopter. The mass of air is further accelerated by the idealized constraints of the flow to

twice the induced velocity at about the distance of one rotor diameter (Figure 3-15).

Figure 3-15 Induced Velocity Idealized for Momentum Theory

Induced power is that portion of total power used to accelerate air downward and create lift. In

equilibrium, the thrust or force generated by the rotor must equal aircraft weight, so the induced

power required to hover is then a function of aircraft weight. If the helicopter weighs more, it

requires more torque, higher rotor speed, less induced velocity, or some combination of all three.

313. BLADE ELEMENT THEORY

Whereas the Momentum Theory can describe the overall forces on the entire rotor disk, the

Blade Element Theory allows for a greater fidelity in understanding the action and reaction of

individual blades within a rotor disk. The basis of Blade Element Theory is to take a very small

portion of the rotor blade and determine the forces acting on it.

Observe airflow conditions at the blade element in Figure 3-17. The blade “sees” a combination

of linear velocity (sometimes called linear flow) and downward induced velocity as

components of relative wind.

The angle of attack (AOA) is the aerodynamic angle formed between the relative wind and the

chord line (Figure 3-18). The pitch angle is the mechanical angle formed between the tip path

plane (TPP) and the chord line. Lift, which is the component of the total aerodynamic force

perpendicular to the relative wind, is tilted aft. This rearward component generated by lift is

induced drag, formed from the acceleration of a mass of air (downwash) and the energy spent in

the creation of trailing vortices. The remaining arrow labeled profile drag is the result of air



friction acting on the blade element and comprised of viscous drag (skin friction) and form drag,

which is the drag produced from the low velocity/low static pressure air formed in the wake of

each blade.

Figure 3-16 Variables in the Blade Element Theory

Figure 3-17 Airflow over the Airfoil



Figure 3-18 Blade Element Diagram



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AIRFOILS 4-1

CHAPTER FOUR

AIRFOILS

400. INTRODUCTION

The purpose of this chapter is to aid the student in understanding airfoils, terminology, and

airflow and associated reactions in the rotor disk.









402. AIRFOIL

Helicopters are able to fly due to aerodynamic forces produced when air passes around the

airfoil. An airfoil is any surface producing more lift than drag when passing through the air at a

suitable angle. Airfoils are most often associated with production of lift. Airfoils are also used

for stability (fin), control (elevator), and thrust or propulsion (propeller or rotor). Certain

airfoils, such as rotor blades, combine some of these functions. The main and tail rotor blades of

the helicopter are airfoils, and air is forced to pass around the blades by mechanically powered

rotation. In some conditions, parts of the fuselage, such as the vertical and horizontal stabilizers,

can become airfoils. Airfoils are carefully structured to accommodate a specific set of flight

characteristics.

403. AIRFOIL TERMINOLOGY AND DEFINITIONS

Blade span: the length of the rotor blade from center of rotation to tip of the blade.

Chord line: a straight line intersecting leading and trailing edges of the airfoil.

Chord: the length of the chord line from leading edge to trailing edge; it is the characteristic

longitudinal dimension of the airfoil section.

CHAPTER FOUR HELICOPTER AERODYNAMICS WORKBOOK

4-2 AIRFOILS

Figure 4-1 Aerodynamic terms of an airfoil

Mean camber line: a line drawn halfway between the upper and lower surfaces of the airfoil.

The chord line connects the ends of the mean camber line. Camber refers to curvature of the

airfoil and subsequent curvature of the mean camber line. The shape of the mean camber is

important for determining aerodynamic characteristics of an airfoil section. Maximum camber

(displacement of the mean camber line from the chord line) and its location help to define the

shape of the mean camber line. The location of maximum camber and its displacement from the

chord line are expressed as fractions or percentages of the basic chord length. By varying the

point of maximum camber, the manufacturer can tailor an airfoil for a specific purpose. The

profile thickness and thickness distribution are important properties of an airfoil section.

Leading edge: the front edge of an airfoil.

Flightpath velocity: the speed and direction of the airfoil passing through the air. For airfoils

on an airplane, the flightpath velocity is equal to true airspeed (TAS). For helicopter rotor

blades, flightpath velocity is equal to rotational velocity, plus or minus a component of

directional airspeed.

Rotational Velocity: Rotational velocity of the blade is constant along the span of the blade.

Rotational velocity is RPM or angular velocity. It does not change. The linear velocity of the

blade element and the relative velocity of the blade element changes and this can be defined as.

rotational relative wind, or the relative velocity of the blade.

Relative wind: defined as the airflow relative to an airfoil and is created by movement of an

airfoil through the air. This is rotational relative wind for rotary-wing aircraft and is covered in

detail later. As an induced airflow may modify flightpath velocity, relative wind experienced by

the airfoil may not be exactly opposite its direction of travel.

Trailing edge: the rearmost edge of an airfoil.

Induced flow: the downward flow of air through the rotor disk.

Resultant relative wind: relative wind modified by induced flow.

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER FOUR

AIRFOILS 4-3

AOA: the angle measured between the resultant relative wind and chord line.

Angle of incidence (AOI): the angle between the chord line of a blade and rotor hub,

commonly referred to as blade pitch angle. For fixed airfoils, such as vertical fins or elevators,

angle of incidence is the angle between the chord line of the airfoil and a selected reference plane

of the helicopter.

Center of pressure: the point along the chord line of an airfoil through which all aerodynamic

forces are considered to act. Since pressures vary on the surface of an airfoil, an average

location of pressure variation is needed. As the AOA changes, these pressures change, and the

center of pressure moves along the chord line.

404. AIRFOIL TYPES

Symmetrical Airfoil

The symmetrical airfoil is distinguished by having identical upper and lower surfaces. The mean

camber line and chord line are the same on a symmetrical airfoil, and it produces no lift at zero

AOA. Most light helicopters incorporate symmetrical airfoils in the main rotor blades.

Nonsymmetrical Airfoil (Cambered)

The nonsymmetrical airfoil has different upper and lower surfaces, with a greater curvature of

the airfoil above the chord line than below. The mean camber line and chord line are different.

The nonsymmetrical airfoil design can produce useful lift at zero AOA. A nonsymmetrical

design has advantages and disadvantages. The advantages are more lift production at a given

AOA than a symmetrical design, an improved lift-to-drag ratio, and better stall characteristics.

The disadvantages are center of pressure travel of up to 20 percent of the chord line (creating

undesirable torque on the airfoil structure) and greater production costs.

Figure 4-2 Airfoil Types


4-4 AIRFOILS

405. ROTOR AXIS

Because of lift differential due to differing rotational relative wind values along the blade, the

blade should be designed with a twist to alleviate internal blade stress and distribute the lifting

force more evenly along the blade. Blade twist provides higher pitch angles at the root where

velocity is low and lower pitch angles nearer the tip where velocity is higher. This increases the

induced air velocity and blade loading near the inboard section of the blade.

Figure 4-3 Blade Twist

Rotor Blade and Hub Definitions

Hub: on the mast, the attaching point for the root of the blade, and the axis about which the

blades rotate.

Tip: the farthest outboard section of the rotor blade

Root: the inner end of the blade and is the point that attaches to the hub

Twist: the change in blade incidence from the root to the outer blade

The angular position of the main rotor blades (as viewed from above, as they rotate about the

vertical axis of the mast) is measured from the helicopter’s longitudinal axis, and usually from its

nose. The radial position of a segment of the blade is the distance from the hub as a fraction of

the total distance.

406. AIRFLOW AND REACTIONS IN THE ROTOR DISK

1. Relative Wind

Knowledge of relative wind is essential for an understanding of aerodynamics and its practical

flight application for the pilot. Relative wind is airflow relative to an airfoil. Movement of an

airfoil through the air creates relative wind. Relative wind moves in a direction parallel to but

opposite of the movement of the airfoil. There are two parts to wind passing a rotor blade:

Horizontal part: caused by the blades turning plus movement of the helicopter through the air


AIRFOILS 4-5

Vertical part: caused by the air being forced down through the rotor blades plus any movement

of the air relative to the blades caused by the helicopter climbing or descending

Figure 4-4 Relative Wind

Figure 4-5 Horizontal component of relative wind

Figure 4-6 Induced Flow


4-6 AIRFOILS

Figure 4-7 Spanwise induced flow velocities

2. Rotational Relative Wind (Tip-Path Plane)

The rotation of rotor blades as they turn about the mast produces rotational relative wind

(tip-path plane). The term rotational refers to the method of producing relative wind. Rotational

relative wind flows opposite the physical flightpath of the airfoil, striking the blade at 90° to the

leading edge and parallel to the plane of rotation; and it is constantly changing in direction

during rotation. Rotational relative wind velocity is highest at blade tips, decreasing uniformly

to zero at the axis of rotation (center of the mast).

Figure 4-8 Rotational Relative Wind

3. Resultant Relative Wind

The resultant relative wind at a hover is rotational relative wind modified by induced flow. This

is inclined downward at some angle and opposite the effective flightpath of the airfoil, rather

than the physical flightpath (rotational relative wind). The resultant relative wind also serves as

the reference plane for development of lift, drag, and total aerodynamic force (TAF) vectors on

the airfoil. When the helicopter has horizontal motion, airspeed further modifies the resultant

relative wind. The airspeed component of relative wind results from the helicopter moving

through the air. This airspeed component is added to, or subtracted from, the rotational relative

wind depending on whether the blade is advancing or retreating in relation to helicopter

movement. Introduction of airspeed relative wind also modifies induced flow. Generally, the

downward velocity of induced flow is reduced. The pattern of air circulation through the disk

changes when the aircraft has horizontal motion. As the helicopter gains airspeed, the addition

of forward velocity results in decreased induced flow velocity. This change results in an

improved efficiency (additional lift) being produced from a given blade pitch setting.


AIRFOILS 4-7

Figure 4-9 Resultant relative wind

Induced Flow (Downwash): At flat pitch, air leaves the trailing edge of the rotor blade in the

same direction it moved across the leading edge; no lift or induced flow is being produced. As

blade pitch angle is increased, the rotor disk induces a downward flow of air through the rotor

blades creating a downward component of air that is added to the rotational relative wind.

Because the blades are moving horizontally, some of the air is displaced downward. The blades

travel along the same path and pass a given point in rapid succession. Rotor blade action

changes the still air to a column of descending air. Therefore, each blade has a decreased AOA

due to the downwash. This downward flow of air is called induced flow (downwash). It is most

pronounced at a hover under no-wind conditions.

Figure 4-10 Induced flow in forward flight

In Ground Effect (IGE): Ground effect is the increased efficiency of the rotor disk caused by

interference of the airflow when near the ground. The air pressure or density is increased, which

acts to decrease the downward velocity of air. Ground effect permits relative wind to be more

horizontal, lift vector to be more vertical, and induced drag to be reduced. These conditions


4-8 AIRFOILS

allow the rotor disk to be more efficient. Maximum ground effect is achieved when hovering

over smooth hard surfaces. When hovering over surfaces as tall grass, trees, bushes, rough

terrain, and water, maximum ground effect is reduced. Rotor efficiency is increased by ground

effect to a height of about one rotor diameter (measured from the ground to the rotor disk) for

most helicopters. Since the induced flow velocities are decreased, the AOA is increased, which

requires a reduced blade pitch angle and a reduction in induced drag. This reduces the power

required to hover IGE.

Figure 4-11 In Ground Effect (IGE)

Out of Ground Effect (OGE): The benefit of placing the helicopter near the ground is lost

above IGE altitude. Above this altitude, the power required to hover remains nearly constant,

given similar conditions (such as wind). Induced flow velocity is increased, resulting in a

decrease in AOA and a decrease in lift. Under the correct circumstances, this downward flow

can become so localized that the helicopter and locally disturbed air will sink at alarming rates.

This effect is called vortex ring state (formerly referenced as settling-with-power) and is

discussed later. A higher blade pitch angle is required to maintain the same AOA as in IGE

hover. The increased pitch angle also creates more drag. This increased pitch angle and drag

requires more power to hover OGE than IGE.


AIRFOILS 4-9

Figure 4-12 Out of Ground Effect (OGE)

407. ROTOR BLADE ANGLES

There are two angles that enable a rotor disk to produce the lift required for a helicopter to fly:

angle of incidence and angle of attack.

Angle of Incidence: Angle of incidence is the angle between the chord line of a main or tail

rotor blade and its rotor disk. It is a mechanical angle rather than an aerodynamic angle and is

sometimes referred to as blade pitch angle. In the absence of induced flow, AOA and angle of

incidence are the same. Whenever induced flow, up flow (inflow), or airspeed modifies the

relative wind, the AOA is different from the angle of incidence. Collective input and cyclic

feathering change the angle of incidence. A change in the angle of incidence changes the AOA,

which changes the coefficient of lift, thereby changing the lift produced by the airfoil.

Figure 4-13 Angle of Incidence


4-10 AIRFOILS

Angle of Attack: AOA is the angle between the airfoil chord line and resultant relative wind. It

is an aerodynamic angle and not easy to measure. It can change with no change in the blade

pitch angle (angle of incidence, discussed earlier).

Figure 4-14 Angle of Attack

When the AOA is increased, air flowing over the airfoil is diverted over a greater distance,

resulting in an increase of air velocity and more lift. As the AOA is increased further, it becomes

more difficult for air to flow smoothly across the top of the airfoil. At this point, the airflow

begins to separate from the airfoil and enters a burbling or turbulent pattern. The turbulence

results in a large increase in drag and loss of lift in the area where it is taking place. Increasing

the AOA increases lift until the critical angle of attack is reached. Any increase in the AOA

beyond this point produces a stall and a rapid decrease in lift.

Several factors may change the rotor blade AOA. The pilot has little direct control over AOA

except indirectly through the flight control input. Collective and cyclic feathering help to make

these changes. Feathering is the rotation of the blade about its longitudinal axis by

collective/cyclic inputs causing changes in blade pitch angle. Collective feathering changes

angle of incidence equally and in the same direction on all rotor blades simultaneously. This

action changes AOA, which changes coefficient of lift (CL), and affects overall lift of the rotor

disk.

Cyclic feathering changes the blade’s AOA differentially around the rotor disk and creates a

differential lift. Aviators use cyclic feathering to control attitude of the rotor disk. It is the

means to control rearward tilt of the rotor (blowback) caused by flapping action and (along with

blade flapping) counteract dissymmetry of lift (discussed later). Cyclic feathering causes attitude

of the rotor disk to change but does not change the amount of net lift the rotor disk is producing.

Most of the changes in AOA come from change in airspeed and rate of climb or descent; others

such as flapping occur automatically due to the rotor system design. Flapping is the up and

down movement of rotor blades about a hinge on a fully articulated rotor system. A semi-rigid

system does not have a hinge but flap as a unit. A rigid rotor system has no vertical or horizontal

hinges, so the blades cannot flap or drag, but they can flex. By flexing, the blades themselves

compensate for the forces which previously required rugged hinges. It occurs in response to

changes in lift due to changing velocity or cyclic feathering. No flapping occurs when the

tip-path plane is perpendicular to the mast. The flapping action alone, or along with cyclic


AIRFOILS 4-11

feathering, controls dissymmetry of lift. Flapping is the primary means of compensating for

dissymmetry of lift.

Pilots adjust AOA through normal control manipulation of the pitch angle of the blades. If the

pitch angle is increased, the AOA increases; if the pitch angle is reduced, the AOA is reduced.


4-12 AIRFOILS

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POWERED FLIGHT 5-1

CHAPTER FIVE

POWERED FLIGHT

500. INTRODUCTION

This chapter introduces powered flight analysis and associated effects.









502. POWERED FLIGHT

In powered flight (hovering, vertical, forward, sideward, or rearward), the total lift and thrust

forces of a rotor are perpendicular to the rotor disk.

503. HOVERING FLIGHT

1. General

Hovering is the most challenging part of flying a helicopter. This is because a helicopter

generates its own gusty air while in a hover, which acts against the fuselage and flight control

surfaces. The end result is constant control inputs and corrections by the pilot to keep the

helicopter where it is required to be. Despite the complexity of the task, the control inputs in a

hover are simple. The cyclic is used to eliminate drift in the horizontal plane, controlling

forward, backward, right and left movement or travel. The throttle, if not governor controlled, is

used to control revolutions per minute (RPM). The collective is used to maintain altitude. The

pedals are used to control nose direction or heading. It is the interaction of these controls that

makes hovering difficult, since an adjustment in any one control requires an adjustment of the

other two, creating a cycle of constant correction. During hovering flight, a helicopter maintains

a constant position over a selected point, usually a few feet above the ground. The ability of the

helicopter to hover comes from the both the lift component, which is the force developed by the

main rotor(s) to overcome gravity and aircraft weight, and the thrust component, which acts

horizontally to accelerate or decelerate the helicopter in the desired direction. Pilots direct the

CHAPTER FIVE HELICOPTER AERODYNAMICS WORKBOOK

5-2 POWERED FLIGHT

thrust of the rotor disk by using the cyclic to rotate the rotor disk plane relative to the horizon.

They do this in order to induce travel or compensate for the wind and hold a position. At a hover

in a no-wind condition, all opposing forces (lift, thrust, drag, and weight) are in balance; they are

equal and opposite. Therefore, lift and weight are equal, resulting in the helicopter remaining at

a stationary hover.

While hovering, the amount of main rotor thrust can be adjusted to maintain the desired hovering

height. This is done by changing the angle of incidence (by moving the collective) of the rotor

blades, and hence their AOA. Changing the AOA changes the drag on the rotor blades, and the

power delivered by the engine must change as well to keep the rotor speed constant (The Nf

governor in the TH-57 works to accomplish this resultant constant RPM, or Nr).

The weight that must be supported is the total weight of the helicopter and its occupants. If the

amount of lift is greater than the actual weight, the helicopter accelerates upwards until the lift

force equals the weight of the helicopter; if lift is less than weight, the helicopter accelerates

downward.

The drag of a hovering helicopter is mainly induced drag incurred while the blades are producing

lift. There is, however, some profile drag on the blades as they rotate through the air and a small

amount of parasite drag from the non-lift-producing surfaces of the helicopter, such as the rotor

hub, cowlings, and landing gear. Throughout the rest of this discussion, the term “drag” includes

induced, profile and parasite drag.

An important consequence of producing thrust is torque. As discussed earlier, Newton’s Third

Law states: for every action there is an equal and opposite reaction. Therefore, as the engine

turns the main rotor disk in a counterclockwise direction, the helicopter fuselage wants to turn

clockwise. The amount of torque is directly related to the amount of engine power being used to

turn the main rotor disk. As power changes, torque changes.

To counteract this torque-induced turning tendency, an anti-torque rotor or tail rotor is

incorporated into most helicopter designs. A pilot can vary the amount of thrust produced by the

tail rotor in relation to the amount of torque produced by the engine. As the engine supplies

more power to the main rotor, the tail rotor must produce more thrust to overcome the increased

torque effect. This control change is accomplished through the use of anti-torque pedals.

2. Translating Tendency (Drift)

During hovering flight, a single main rotor helicopter tends to move in the direction of tail rotor

thrust. This lateral (or sideward) movement is called translating tendency.

To counteract this tendency, one or more of the following features may be used. All examples

are for a counterclockwise rotating main rotor disk.

a. The main transmission is mounted at a slight angle to the left (when viewed from

behind) so that the rotor mast has a built-in tilt to oppose the tail rotor thrust.

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER FIVE

POWERED FLIGHT 5-3

b. Flight controls can be rigged so that the rotor disk is tilted to the left slightly when the

cyclic is centered. Whichever method is used, the tip-path plane is tilted slightly to

the left in the hover.

c. The transmission is mounted so the rotor shaft is vertical with respect to the fuselage,

the helicopter “hangs” left skid low in the hover. (The opposite is true for rotor disks

turning clockwise when viewed from above.)

d. The helicopter fuselage will also be tilted when the tail rotor is below the main rotor

disk and supplying anti-torque thrust. The fuselage tilt is caused by the imperfect

balance of the tail rotor thrust against the main rotor torque in the same plane. The

helicopter tilts due to two separate forces, the main rotor disk tilt to neutralize the

translating tendency and the lower tail rotor thrust below the plane of the torque

action.

e. In forward flight, the tail rotor continues to push to the right, and the helicopter makes

a small angle with the wind when the rotors are level and the slip ball is in the middle.

This is called inherent sideslip. For some larger helicopters, the vertical fin or

stabilizer is often designed with the tail rotor mounted on them to correct this side slip

and to eliminate some of the tilting at a hover. (By mounting the tail rotor on top of

the vertical fin or pylon, the anti-torque is more in line with or closer to the horizontal

plane of torque, resulting in less airframe (or body) lean from the tail rotor.) Also,

having the tail rotor higher off the ground reduces the risk of objects coming in

contact with the blades, but at the cost of increased weight and complexity.

Figure 5-1 Translating Tendency


5-4 POWERED FLIGHT

3. Pendular Action

Since the fuselage of the helicopter, with a single main rotor, is suspended from a single point

and has considerable mass, it is free to oscillate either longitudinally or laterally in the same way

as a pendulum. This pendular action can be exaggerated by over controlling; therefore, control

movements should be smooth and not exaggerated.

Figure 5-2 Pendular action

The horizontal stabilizer helps to level the helicopter in forward flight. However, in rearward

flight, the horizontal stabilizer can press the tail downward, resulting in a tail strike if the

helicopter is moved rearward into the wind. Normally, with the helicopter mostly into the wind,

the horizontal stabilizer experiences less headwind component as the helicopter begins rearward

travel (downwind). When rearward flight groundspeed equals the wind speed, then the

helicopter is merely hovering in a no-wind condition. However, rearward hovering into the wind

requires considerable care and caution to prevent tail strikes.

It is important to note that there is a difference in the amount of pendular action between a semi

rigid system and a fully articulated system. Because of the hard connection (offset) of the latter,

the centrifugal force pulling out on the blades is transferred to the fuselage, and the fuselage

tends to follow the rotor attitude. The semi rigid system is a true pendulum, with thrust required

to create a moment around the fuselage CG to allow for control of the fuselage.

4. Coning

In order for a helicopter to generate lift, the rotor blades must be turning. Rotor disk rotation

drives the blades into the air, creating a relative wind component without having to move the


POWERED FLIGHT 5-5

airframe through the air as with an airplane or glider. Depending on the motion of the blades and

helicopter airframe, many factors cause the relative wind direction to vary. The rotation of the

rotor disk creates centrifugal force (inertia), which tends to pull the blades straight outward from

the main rotor hub: the faster the rotation, the greater the centrifugal force, the slower the

rotation, the smaller the centrifugal force. This force gives the rotor blades their rigidity and, in

turn, the strength to support the weight of the helicopter. The maximum centrifugal force

generated is determined by the maximum operating rotor revolutions per minute (RPM).

As lift on the blades is increased (in a takeoff, for example), two major forces are acting at the

same time—centrifugal force acting outward, and lift acting upward. The result of these two

forces is that the blades assume a conical path instead of remaining in the plane perpendicular to

the mast. This can be seen in any helicopter when it takes off; the rotor disk changes from flat to

a slight cone shape.

Figure 5-3 Coning

If the rotor RPM is allowed to go too low (below the minimum power-on rotor RPM, for

example), the centrifugal force becomes smaller and the coning angle becomes much larger. In

other words, should the RPM decrease too much, at some point the rotor blades fold up with no

chance of recovery.

5. Coriolis Effect

The Coriolis Effect is also referred to as the law of conservation of angular momentum. It states

that the value of angular momentum of a rotating body does not change unless an external force

is applied. In other words, a rotating body continues to rotate with the same rotational velocity

until some external force is applied to change the speed of rotation. Angular momentum is the

moment of inertia (mass times distance from the center of rotation squared) multiplied by the

speed of rotation.

Changes in angular velocity, known as angular acceleration and deceleration, take place as the

mass of a rotating body is moved closer to or farther away from the axis of rotation. The speed

of the rotating mass varies proportionately with the square of the radius.


5-6 POWERED FLIGHT

An excellent example of this principle in action is a figure skater performing a spin on ice skates.

The skater begins rotation on one foot, with the other leg and both arms extended. The rotation

of the skater’s body is relatively slow. When a skater draws both arms and one leg inward, the

moment of inertia (mass times radius squared) becomes much smaller and the body is rotating

almost faster than the eye can follow. Because the angular momentum must, by law of nature,

remain the same (no external force applied), the angular velocity must increase.

The rotor blade rotating about the rotor hub possesses angular momentum. As the rotor begins to

cone due to G-loading maneuvers, the diameter of the rotor disk shrinks. Due to conservation of

angular momentum, the blades increase speed even though the blade tips have a shorter distance

to travel due to reduced disk diameter. The action results in an increase in rotor RPM which

causes a slight increase in lift. Most pilots arrest this increase of RPM with an increase in

collective pitch. This increase in blade RPM lift is somewhat negated by the slightly smaller

disk area as the blades cone upward.

6. Gyroscopic Precession

The spinning main rotor of a helicopter acts like a gyroscope. As such, it has the properties of

gyroscopic action, one of which is precession. Gyroscopic precession is the resultant action or

deflection of a spinning object when a force is applied to this object. This action occurs

approximately 90° in the direction of rotation from the point where the force is applied

(or 90° later in the rotation cycle).

Figure 5-4 Gyroscopic precession

Examine a two-bladed rotor disk to see how gyroscopic precession affects the movement of the

tip-path plane. Moving the cyclic pitch control increases the angle of incidence of one rotor

blade with the result of a greater lifting force being applied at that point in the plane of rotation.


POWERED FLIGHT 5-7

This same control movement simultaneously decreases the angle of incidence of the other blade

the same amount, thus decreasing the lifting force applied at that point in the plane of rotation.

The blade with the increased angle of incidence tends to flap up; the blade with the decreased

angle of incidence tends to flap down. Because the rotor disk acts like a gyro, the blades reach

maximum deflection at a point approximately 90° later in the plane of rotation. Figure 5-5

illustrates the result of a forward cyclic input. The retreating blade angle of incidence is

increased, and the advancing blade angle of incidence is decreased resulting in a tipping forward

of the tip-path plane, since maximum deflection takes place 90° later when the blades are at the

rear and front, respectively.

Figure 5-5 Forward Cyclic Input

In a rotor disk using three or more blades, the movement of the cyclic pitch control changes the

angle of incidence of each blade an appropriate amount so that the end result is the same.

504. VERTICAL FLIGHT

Hovering is actually an element of vertical flight. Increasing the angle of incidence of the rotor

blades (pitch) while keeping their rotation speed constant generates additional lift and the

helicopter ascends. Decreasing the pitch causes the helicopter to descend. In a no-wind

condition in which lift and thrust are less than weight and drag, the helicopter descends

vertically. If lift and thrust are greater than weight and drag, the helicopter ascends vertically.


5-8 POWERED FLIGHT

Figure 5-6 No wind hover

505. FORWARD FLIGHT

In steady forward flight, with no change in airspeed or vertical speed, the four forces of lift,

thrust, drag, and weight must be in balance. Once the tip-path plane is tilted forward, the total

lift-thrust force is also tilted forward. This resultant lift-thrust force can be resolved into two

components—lift acting vertically upward and thrust acting horizontally in the direction of

flight. In addition to lift and thrust, there is weight (the downward acting force) and drag

(the force opposing the motion of an airfoil through the air).

Figure 5-7 Transition to forward flight


POWERED FLIGHT 5-9

In straight-and-level, un-accelerated forward flight (straight-and-level flight is flight with a

constant heading and at a constant altitude), lift equals weight and thrust equals drag. If lift

exceeds weight, the helicopter accelerates vertically until the forces are in balance; if thrust is

less than drag, the helicopter slows down until the forces are in balance. As a helicopter initiates

a move forward, it begins to lose altitude because lift is lost as thrust is diverted forward.

However, as the helicopter begins to accelerate from a hover, the rotor disk becomes more

efficient due to translational lift. The result is excess power over that which is required to hover.

Continued acceleration causes an even larger increase in airflow through the rotor disk (up to a

maximum determined by drag and the engine’s limit of power), and more efficient flight. In

order to maintain un-accelerated flight, the pilot must understand that with any changes in power

or in cyclic movement, the helicopter begins either to climb or to descend. Once straight-and-

level flight is obtained, the pilot should make note of the power (torque setting) required and not

make major adjustments to the flight controls.

Figure 5-8 Power vs. airspeed chart

1. Airflow in forward flight

Airflow across the rotor disk in forward flight varies from airflow at a hover. In forward flight,

air flows opposite the aircraft’s flightpath. The velocity of this air flow equals the helicopter’s

forward speed. Because the rotor blades turn in a circular pattern, the velocity of airflow across

a blade depends on the position of the blade in the plane of rotation at a given instant, its

rotational velocity, and airspeed of the helicopter. Therefore, the airflow meeting each blade

varies continuously as the blade rotates. The highest velocity of airflow occurs over the right

side (3 o’clock position) of the helicopter (advancing blade in a rotor disk that turns

counterclockwise) and decreases to rotational velocity over the nose. It continues to decrease

until the lowest velocity of airflow occurs over the left side (9 o’clock position) of the helicopter

(retreating blade). As the blade continues to rotate, velocity of the airflow then increases to

rotational velocity over the tail. It continues to increase until the blade is back at the 3 o’clock

position.


5-10 POWERED FLIGHT

The advancing blade in Figure 5-9 position A moves in the same direction as the helicopter. The

velocity of the air meeting this blade equals rotational velocity of the blade plus wind velocity

resulting from forward airspeed. The retreating blade (position C) moves in a flow of air moving

in the opposite direction of the helicopter. The velocity of airflow meeting this blade equals

rotational velocity of the blade minus wind velocity resulting from forward airspeed. The blades

(positions B and D) over the nose and tail move essentially at right angles to the airflow created

by forward airspeed; the velocity of airflow meeting these blades equals the rotational velocity.

This results in a change to velocity of airflow all across the rotor disk and a change to the lift

pattern of the rotor disk.

Figure 5-9 Airflow in forward flight

Advancing Blade: As the relative wind speed of the advancing blade increases, the blade gains

lift and begins to flap up. It reaches its maximum up-flap velocity at the 3 o’clock position,

where the wind velocity is the greatest. This up-flap creates a downward flow of air and has the

same effect as increasing the induced flow velocity by imposing a downward vertical velocity

vector to the relative wind which decreases the AOA.


POWERED FLIGHT 5-11

Retreating Blade: As relative wind speed of the retreating blade decreases, the blade loses lift

and begins to flap down. It reaches its maximum down-flap velocity at the 9 o’clock position,

where wind velocity is the least. This down-flap creates an upward flow of air and has the same

effect as decreasing the induced flow velocity by imposing an upward velocity vertical vector to

the relative wind which increases the AOA.

Dissymmetry of Lift: Dissymmetry of lift is the differential (unequal) lift between advancing

and retreating halves of the rotor disk caused by the different wind flow velocity across each

half. This difference in lift would cause the helicopter to be uncontrollable in any situation other

than hovering in a calm wind. There must be a means of compensating, correcting, or

eliminating this unequal lift to attain symmetry of lift.

When the helicopter moves through the air, the relative airflow through the main rotor disk is

different on the advancing side from the retreating side. The relative wind encountered by the

advancing blade is increased by the forward speed of the helicopter, while the relative wind

speed acting on the retreating blade is reduced by the helicopter’s forward airspeed. Therefore,

as a result of the relative wind speed, the advancing blade side of the rotor disk can produce

more lift than the retreating blade side.

Figure 5-10 Dissymmetry of lift

If this condition were allowed to exist, a helicopter with a counterclockwise main rotor blade

rotation would roll to the left because of the difference in lift. In reality, the main rotor blades

flap and feather automatically to equalize lift across the rotor disk. Articulated rotor disks,

usually with three or more blades, incorporate a horizontal hinge (flapping hinge) to allow the

individual rotor blades to move, or flap up and down as they rotate. A semi-rigid rotor disk

(two blades) utilizes a teetering hinge, which allows the blades to flap as a unit. When one blade

flaps up, the other blade flaps down.


5-12 POWERED FLIGHT

As shown in Figure 5-11, as the rotor blade reaches the advancing side of the rotor disk (A), it

reaches its maximum up flap velocity. When the blade flaps upward, the angle between the

chord line and the resultant relative wind decreases. This decreases the AOA, which reduces the

amount of lift produced by the blade. At position (C), the rotor blade is now at its maximum

down flapping velocity. Due to down flapping, the angle between the chord line and the

resultant relative wind increases. This increases the AOA and thus the amount of lift produced

by the blade.

Figure 5-11 Effect of flapping

The combination of blade flapping and slow relative wind acting on the retreating blade normally

limits the maximum forward speed of a helicopter. At a high forward speed, the retreating blade

stalls because of a high AOA and slow relative wind speed. This situation is called retreating

blade stall and is evidenced by a nose pitch up, vibration, and a rolling tendency, usually to the

left in helicopters with counterclockwise blade rotation.

Pilots can avoid retreating blade stall by not exceeding the never-exceed speed. This speed is

designated VNE and is indicated on a placard and marked on the airspeed indicator by a red line.

Blade flapping compensates for dissymmetry of lift in the following way. At a hover, equal lift

is produced around the rotor disk with equal pitch (AOI) on all the blades and at all points in the

rotor disk (disregarding compensation for translating tendency). The rotor disk is parallel to the

horizon. To develop a thrust force, the rotor disk must be tilted in the desired direction of

movement. Cyclic feathering changes the angle of incidence differentially around the rotor disk.

For a counterclockwise rotation, forward cyclic movement decreases the angle of incidence on

the right of the rotor disk and increases it on the left.


POWERED FLIGHT 5-13

When transitioning to forward flight either from a hover or taking off from the ground, pilots

must be aware that as the helicopter speed increases, translational lift becomes more effective

and causes the nose to rise or pitch up (sometimes referred to as blowback). This tendency is

caused by the combined effects of dissymmetry of lift and transverse flow. Pilots must correct

for this tendency by maintaining a constant rotor disk attitude that will move the helicopter

through the speed range in which blowback occurs. If the nose is permitted to pitch up while

passing through this speed range, the aircraft may also tend to roll to the right. To correct for

this tendency, the pilot must continuously move the cyclic forward as velocity of the helicopter

increases until the takeoff is complete, and the helicopter has transitioned into forward flight.

Figure 5-12 illustrates the tilting forward of the rotor disk, which is the result of a change in pitch

angle with forward cyclic. At a hover, the cyclic is centered and the pitch angle on the

advancing and retreating blades is the same. At low forward speeds, moving the cyclic forward

reduces pitch angle on the advancing blade and increases pitch angle on the retreating blade.

This causes a slight rotor disk tilt. At higher forward speeds, the pilot must continue to move the

cyclic forward. This further reduces pitch angle on the advancing blade and further increases

pitch angle on the retreating blade. As a result, there is even more tilt to the rotor disk than at

lower speeds.

Figure 5-12 Blowback

A horizontal lift component (thrust) generates higher helicopter airspeed. The higher airspeed

induces blade flapping to maintain symmetry of lift. The combination of flapping and cyclic

feathering maintains symmetry of lift and desired attitude on the rotor disk and helicopter.

2. Translational lift

Improved rotor efficiency resulting from directional flight is called translational lift. The

efficiency of the hovering rotor disk is greatly improved with each knot of incoming wind gained

by horizontal movement of the aircraft or surface wind. As the incoming wind produced by

aircraft movement or surface wind enters the rotor disk, turbulence and vortices are left behind

and the flow of air becomes more horizontal. In addition, the tail rotor becomes more

aerodynamically efficient during the transition from hover to forward flight.


5-14 POWERED FLIGHT

Figure 5-13 Airflow with minimal headwind

Figure 5-14 Airflow just prior to ETL

Effective Translational Lift (ETL): While transitioning to forward flight at about 16 to 24

knots, the helicopter goes through effective translational lift (ETL). As mentioned earlier in the

discussion on translational lift, the rotor blades become more efficient as forward airspeed

increases. Between 16 and 24 knots, the rotor disk completely outruns the recirculation of old

vortices and begins to work in relatively undisturbed air. The flow of air through the rotor disk

is more horizontal, which reduces induced flow and drag with a corresponding increase in angle

of attach and lift. The additional lift available at this speed is referred to as the ETL, which

makes the rotor disk operate more efficiently. This increased efficiency continues with increased

airspeed until the best climb airspeed is reached, and total drag is at its lowest point.

As speed increases, translational lift becomes more effective, nose rises or pitches up, and

aircraft rolls to the right. The combined effects of dissymmetry of lift, gyroscopic precession,

and transverse flow effect cause this tendency. It is important to understand these effects and

anticipate correcting for them. Once the helicopter is transitioning through ETL, the pilot needs

to apply forward and left lateral cyclic input to maintain a constant rotor-disk attitude.


POWERED FLIGHT 5-15

Figure 5-15 ETL

Translational Thrust: Translational thrust occurs when the tail rotor becomes more

aerodynamically efficient during the transition from hover to forward flight. As the tail rotor

works in progressively less turbulent air, this improved efficiency produces more anti-torque

thrust, causing the nose of the aircraft to yaw left (with a main rotor turning counterclockwise)

and forces the pilot to apply right pedal (decreasing the AOA in the tail rotor blades) in response.

In addition, during this period, the airflow affects the horizontal components of the stabilizer

found on most helicopters which tends to bring the nose of the helicopter to a more level attitude.

3. Induced flow

As the rotor blades rotate, they generate what is called rotational relative wind. This airflow is

characterized as flowing parallel and opposite the rotor’s plane of rotation and striking

perpendicular to the rotor blade’s leading edge. This rotational relative wind is used to generate

lift. As rotor blades produce lift, air is accelerated over the foil and projected downward.

Anytime a helicopter is producing lift, it moves large masses of air vertically and down through

the rotor disk. This downwash or induced flow can significantly change the efficiency of the

rotor disk. Rotational relative wind combines with induced flow to form the resultant relative

wind. As induced flow increases, resultant relative wind becomes less horizontal. Since AOA is

determined by measuring the difference between the chord line and the resultant relative wind, as

the resultant relative wind becomes less horizontal, AOA decreases.

4. Transverse flow effect

As the helicopter accelerates in forward flight, induced flow drops to near zero at the forward

disk area and increases at the aft disk area. These differences in lift between the fore and aft

portions of the rotor disk are called transverse flow effect. This increases the AOA at the front

disk area causing the rotor blade to flap up and reduces AOA at the aft disk area causing the

rotor blade to flap down. Because the rotor acts like a gyro, maximum displacement occurs 90°


5-16 POWERED FLIGHT

in the direction of rotation. The result is a tendency for the helicopter to roll slightly to the right

as it accelerates through approximately 20 knots or if the headwind is approximately 20 knots.

Transverse flow effect is recognized by increased vibrations of the helicopter at airspeeds around

12 to 15 knots and can be produced by forward flight or from the wind while in a hover. This

vibration happens at an airspeed just below ETL on takeoff and after passing through ETL

during landing. The vibration happens close to the same airspeed as ETL because that is when

the greatest lift differential exists between the front and rear portions of the rotor system. As

such, some pilots confuse the vibration felt by transverse flow effect with passing through ETL.

To counteract transverse flow effect, a cyclic input to the left may be needed.

506. SIDEWARD FLIGHT

In sideward flight, the tip-path plane is tilted in the direction that flight is desired. This tilts the

total lift-thrust vector sideward. In this case, the vertical or lift component is still straight up and

weight straight down, but the horizontal or thrust component now acts sideward with drag acting

to the opposite side.

Figure 5-16 Sideward Flight

Sideward flight can be a very unstable condition due to the parasitic drag of the fuselage

combined with the lack of horizontal stabilizer for that direction of flight. Increased altitudes

help with control and the pilot must always scan in the direction of flight. Movement of the

cyclic in the intended direction of flight causes the helicopter to move, controls the rate of speed,

and ground track, but the collective and pedals are key to successful sideward flight. Just as in

forward flight, the collective keeps the helicopter from contacting the ground and the pedals help

maintain the correct heading; even in sideward flight, the tail of the helicopter should remain


POWERED FLIGHT 5-17

behind you. Inputs to the cyclic should be smooth and controlled, and the pilot should always be

aware of the tip-path plane in relation to the ground.

Contacting the ground with the skids during sideward flight will most likely result in a dynamic

rollover event before the pilot has a chance to react. Extreme caution should be used

when maneuvering the helicopter sideways to avoid such hazards from happening.

507. REARWARD FLIGHT

For rearward flight, the tip-path plane is tilted rearward, which, in turn, tilts the lift-thrust vector

rearward. Drag now acts forward with the lift component straight up and weight straight down.

Figure 5-17 Rearward Flight

Pilots must be aware of the hazards of rearward flight. Because of the position of the horizontal

stabilizer, the tail end of the helicopter tends to pitch downward in rearward flight, causing the

probability of hitting the ground to be greater than in forward flight. Another factor to consider

in rearward flight is skid design. Most helicopter skids are not turned upward in the back, and

any contact with the ground during rearward flight can put the helicopter in an uncontrollable

position leading to tail rotor contact with the ground. Pilots must do a thorough scan of the area

before attempting to hover rearward, looking for obstacles and terrain changes. Slower airspeeds

can help mitigate risk and maintain a higher-than-normal hover altitude.


5-18 POWERED FLIGHT

508. TURNING FLIGHT

In forward flight, the rotor disk is tilted forward, which also tilts the total lift-thrust force of the

rotor disk forward. When the helicopter is banked, the rotor disk is tilted sideward resulting in

lift being separated into two components. Lift acting upward and opposing weight is called the

vertical component of lift. Lift acting horizontally and opposing inertia (centrifugal force) is the

horizontal component of lift (centripetal force).

As the angle of bank increases, the total lift force is tilted more toward the horizontal, thus

causing the rate of turn to increase because more lift is acting horizontally. Since the resultant

lifting force acts more horizontally, the effect of lift acting vertically is decreased. To

compensate for this decreased vertical lift, the AOA of the rotor blades must be increased in

order to maintain altitude. The steeper the angle of bank is, the greater the AOA of the rotor

blades required to maintain altitude. Thus, with an increase in bank and a greater AOA, the

resultant lifting force increases, and the rate of turn is higher. Simply put, collective pitch must

be increased in order to maintain altitude and airspeed while turning. Collective pitch controls

the angle of incidence and along with other factors, determines the overall AOA in the rotor disk.

Figure 5-18 Turning Flight

AUTOROTATION 6-1

CHAPTER SIX

AUTOROTATION

600. INTRODUCTION

This chapter introduces the conceptual and aerodynamic principles involved during

autorotational flight.













602. GENERAL

In a helicopter, an autorotative descent is a power-off maneuver in which the engine is

disengaged from the main rotor disk and the rotor blades are driven solely by the upward flow of

air through the rotor. In other words, the engine is no longer supplying power to the main rotor.

At the instant of engine failure, the main rotor blades are producing lift and thrust from their

angle of attack (AOA) and velocity. By lowering the collective (which must be done

immediately in case of an engine failure), lift and drag are reduced, and the helicopter begins an

immediate descent, thus producing an upward flow of air through the rotor disk. This upward

flow of air through the rotor disk provides sufficient thrust to maintain rotor RPM throughout the

descent. The tail rotor is driven by the main rotor transmission during autorotation, so heading

control is maintained with the anti-torque pedals as in normal flight.

CHAPTER SIX HELICOPTER AERODYNAMICS WORKBOOK

6-2 AUTOROTATION

Autorotation is further defined as the state of flight where the main rotor disk of a helicopter is

being turned by the action of air moving up through the rotor rather than engine power driving

the rotor. In normal, powered flight, air is drawn into the main rotor disk from above and

exhausted downward, but during autorotation, air moves up into the rotor disk from below as the

helicopter descends. Autorotation is permitted mechanically by a freewheeling unit: a clutch

mechanism that allows the main rotor to continue turning even if the engine is not running. If

the engine fails, the freewheeling unit automatically disengages the engine from the main rotor

allowing the main rotor to rotate freely. It is the means by which a helicopter can be landed

safely in the event of an engine failure.

Figure 6-1 Airflow in an autorotation

603. VERTICAL AUTOROTATION

Most autorotations are performed with forward speed. For simplicity, the following

aerodynamic explanation describes a vertical autorotative descent (no forward speed) in still air.

Under these conditions, the forces that cause the blades to turn are similar for all blades

regardless of their position in the plane of rotation. Therefore, dissymmetry of lift resulting from

helicopter airspeed is not a factor.

Figure 6-2 Rotor disc regions in autorotation zero speed

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER SIX

AUTOROTATION 6-3

During vertical autorotation, the rotor disk is divided into three regions: driven region, driving

region, and stall region. Figure 6-3 shows three blade sections that illustrate force vectors.

Part A is the driven region, B and D are points of equilibrium, part C is the driving region, and

part E is the stall region. Force vectors are different in each region because rotational relative

wind is slower near the blade root and increases continually toward the blade tip. Also, blade

twist gives a more positive AOA in the driving region than in the driven region. The

combination of the inflow up through the rotor with rotational relative wind produces different

combinations of aerodynamic force at every point along the blade.

The driven region, also called the propeller region, is nearest the blade tips. Normally, it consists

of about 30 percent of the radius. In the driven region, part A of Figure 6-3 the TAF acts behind

the axis of rotation, resulting in an overall drag force. The driven region produces some lift, but

that lift is offset by drag. The overall result is a deceleration in the rotation of the blade. The

size of this region varies with the blade pitch, rate of descent, and rotor RPM. When changing

autorotative RPM blade pitch, or rate of descent, the size of the driven region in relation to the

other regions also changes.

There are two points of equilibrium on the blade, one between the driven region and the driving

region, and one between the driving region and the stall region. At points of equilibrium, TAF is

aligned with the axis of rotation. Lift and drag are produced, but the total effect produces neither

acceleration nor deceleration.

The driving region, or autorotative region, normally lies between 25 to 70 percent of the blade

radius. Part C of Figure 6-3 shows the driving region of the blade, which produces the forces

needed to turn the blades during autorotation. Total aerodynamic force in the driving region is

inclined slightly forward of the axis of rotation, producing a continual acceleration force. This

inclination supplies thrust, which tends to accelerate the rotation of the blade. Driving region

size varies with blade pitch setting, rate of descent, and rotor RPM.

By controlling the size of this region, a pilot can adjust autorotative RPM. For example, if the

collective pitch is raised, the pitch angle increases in all regions. This causes the point of

equilibrium to move inboard along the blade’s span, thus increasing the size of the driven region.

The stall region also becomes larger while the driving region becomes smaller. Reducing the

size of the driving region causes the acceleration force of the driving region and RPM to

decrease. A constant rotor RPM is achieved by adjusting the collective pitch so blade

acceleration forces from the driving region are balanced with the deceleration forces from the

driven and stall regions.

The inner 25 percent of the rotor blade is referred to as the stall region and operates above its

maximum AOA (stall angle), causing drag, which tends to slow rotation of the blade. Part E of

Figure 6-3 depicts the stall region.

604. FORWARD FLIGHT AUTOROTATION

Autorotative force in forward flight is produced in exactly the same manner as when the

helicopter is descending vertically in still air. However, because forward speed changes the


6-4 AUTOROTATION

inflow of air up through the rotor disk, all three regions move outboard along the blade span on

the retreating side of the disk where AOA is larger. With lower AOA on the advancing side

blade, more of the blade falls in the driven region. On the retreating side, more of the blade is in

the stall region. A small section near the root experiences a reversed flow; therefore, the size of

the driven region on the retreating side is reduced.

Prior to landing from an autorotative descent (or autorotation), the pilot must flare the helicopter

in order to decelerate. The pilot initiates the flare by applying aft cyclic. As the helicopter flares

back, the airflow patterns change around the blades causing the RPM to increase. Pilots must

adjust the collective as necessary to keep the RPM within operating limits.

Figure 6-3 Rotor blade regions in autorotation


AUTOROTATION 6-5

Figure 6-4 Rotor disc regions in autorotation forward speed

605. AUTOROTATION DESCENT VARIABLES

Performance in the descent depends upon the forces acting upon the rotor. Airspeed affects the

aerodynamic force on all blades and total power required for flight. Most autorotations are flown

at a 13-17 degree glide angle which offers the lowest descent rate in autorotation. It is worth

noting that very low speed autorotations, while possible, have very high associated descent rates.

Aircraft trim and gross weight also affect power required, so they also affect performance. As an

aircraft moves further out of trim the parasite drag increases, power required increases, and

hence descent rate increases. Gross weight determines RPM at a given collective pitch. At high

gross weight, more blade pitch is required to maintain a desired RPM, so higher gross weights

result in a slower rate of descent, assuming all other variables remain the same. RPM also varies

in descent with altitude (both PA and DA). Higher DA requires higher blade pitch to maintain a

given RPM but, due to the lower air density, a higher rate of descent will still occur.

The factors affecting autorotative descent performance are: Airspeed, Trim, Gross Weight, DA,

and RPM

RPM tradeoffs. RPM is adjusted by varying collective. Adjusting RPM in an autorotation

affects rate of descent and energy stored in the rotor. Selection of a good autorotation rotor

speed depends upon desired performance. High RPM stores energy well and but involves a

higher descent rate. Low RPM provides a slower descent and longer glide, but provides less

stored power for use in the flare. Taken to an extreme, low RPM can stall an excessive portion

of the rotor and make recovery extremely difficult. Specific considerations follow:


6-6 AUTOROTATION

High RPM

1. Centrifugal loads on hub.

2. Excessive propeller region so higher rate of descent.

3. Rotational energy to trade off in a flare.

4. Good for high inertia systems which would have difficulty building RPM rapidly in a flare.

Low RPM

1. Higher AOA therefore a slower rate of descent.

2. Excessive stall region if RPM gets too low resulting in an increase in rate of descent.

3. Less rotational energy to trade off in a flare.

4. Good for low inertia systems which can build RPM rapidly in a flare.

5. Rotor blades lose centrifugal stiffness and cone upwards reducing the effective disk area,

increases material stresses, and increases the rate of descent.

606. PHASES OF AUTOROTATION

Autorotations are divided into three distinct phases: entry, steady-state descent, and deceleration

(flare) and touchdown. Each phase is aerodynamically different from the others.

Level Powered Flight at High Speed. Figure 6-5 shows the airflow and force vectors for a

blade in this configuration. The lift and drag vectors are large, and the total aerodynamic force is

inclined well to the rear of the axis of rotation. An engine failure in this mode will cause rapid

rotor RPM decay. To prevent this, the aviator must lower the collective quickly to reduce drag

and incline the total aerodynamic force vector forward, nearer the axis of rotation.

Figure 6-5 Force Vectors in Level-Powered Flight at High Speed


AUTOROTATION 6-7

Entry. This phase is entered after loss of engine power. The loss of engine power and rotor

RPM is more pronounced when the helicopter is at high gross weight or high forward speed or in

high-DA conditions. Any of these conditions demand increased power and a more abrupt

reaction to the loss of that power. In most helicopters, it takes only seconds for the RPM decay

to fall into a minimum safe range, requiring a quick response from the aviator.

After an engine failure the pilot enters an autorotation by lowering the collective. AOA lessens

as airflow begins to move less downward, then shifts to an upward flow. The net result is that

lower AOA and less pitch make a smaller aerodynamic force that is not tilted as far aft

(Figure 6-6). Vertical force is reduced, so a descent begins, but the associated reduction in drag

keeps the rotor from losing too much RPM.

Collective Pitch Reduction. Figure 6-6 shows the airflow and force vectors for a blade

immediately after power loss and the subsequent collective reduction, yet before the aircraft has

begun to descend. Lift and drag are reduced, with the total aerodynamic force vector inclined

further forward than it was in powered flight. As the helicopter begins to descend, the airflow

begins to flow upward from under the rotor system. This causes the total aerodynamic force to

incline further forward until it reaches an equilibrium that maintains a safe operating RPM

(Figure 6-7).

Figure 6-6 Force Vectors after Power Loss – Reduced Collective

Possible aircraft initial reactions are based upon sudden loss of torque on the main rotor. Thus,

the helicopter will yaw left due to a reduction in anti-torque required (before pedals are

adjusted), and may roll right due to residual tail rotor force. The primary concern should be

controlling RPM. To make the entry a success, blade pitch must be lowered in a timely manner.

The rate of rotor speed decay will determine how quickly the collective must be lowered, and the

rate of decay will be determined by rotor inertia and power required. Rotor inertia is a rotor

head’s resistance to changes in velocity. A high-inertia rotor head will tend to remain at the

same RPM longer after a loss of power or in a flare than a low-inertia rotor head. Power

required is related to induced power, so it is affected by density and airspeed. In practical terms,

this means that the following factors affect successful autorotation entry:

1. Rotor blade pitch (dependent on flight condition - airspeed, gross weight, climb/descent,

etc.)


6-8 AUTOROTATION

2. Rotor inertia

a. High inertia - RPM builds and decays slowly.

b. Low inertia - RPM builds and decays rapidly.

3. Pilot reaction time.

4. Entry altitude (time to establish a stabilized autorotation)

5. Entry airspeed

Steady-State Descent. Figure 6-7 shows the airflow and force vectors for a blade in a

steady-state autorotative descent. Airflow is now upward through the rotor disk because of the

descent. This up-flow of air creates a larger AOA, although blade pitch angle has not changed

since the descent began. Total aerodynamic force on the blade increases and is inclined further

forward until equilibrium is established, rate of descent and rotor RPM stabilize, and the

helicopter is descending at a constant angle. Angle of descent is normally 13 - 17 degrees,

depending on airspeed, DA, wind, and the type of helicopter.

Figure 6-7 Force Vectors in Autorotative Steady-State Descent

When autorotation is established, up-flow tilts the relative wind downward, which moves the net

aerodynamic force forward. In a stabilized autorotation the component of lift in the horizontal

direction balances out the horizontal component of drag so that drag does not reduce the rotor

RPM.


AUTOROTATION 6-9

RPM Stability. In an autorotation, transient changes in aircraft attitude or wind shifts can

change the airflow through the rotor system, therefore affecting RPM. However, the rotor

system demonstrates RPM stability in response to small changes. Figure 6-8 graphically

describes the blade region variations with RPM changes. In autorotation, the blade is at flat pitch

(or near flat pitch), which is designed to maintain a constant RPM at a given descent rate. When

an external force (winds/airflow etc., vice a change in collective setting) causes a small transient

increase in the RPM, the regions shift inboard, enlarging the driven (prop) region and associated

drag, while also reducing the moment arm for the driving (auto) region on the blades. This

reduction in the driving region causes the RPM to decrease back towards the original RPM. Just

the opposite happens with slight decreases in RPM. Thus, for minor RPM variations, the rotor

system has RPM stability.

However, with a large decrease in RPM, even though the driving region of the blade increases,

the stall region and drag it produced also increases. The increased moment arm for the driving

region may not be sufficient to regain the lost RPM before the aircraft reaches the ground.

Figure 6-8 RPM Response to Small RPM Variations

Since the amount of blade surface producing positive autorotative driving force varies according

to RPM and this driving force is synonymous with thrust produced, it is obvious the pilot has

additional control over rate of descent by changing pitch through collective application.

Excessively high Nr produces less driving force and a higher rate of descent, and very low Nr

leads to low driving force in proportion to high drag associated with a stalled profile. There is an

optimum RPM range, which produces the greatest net driving force and minimum descent rate.

It is in the best interest of the pilot to strive for this RPM range until reaching flare altitude. The

pilot must monitor RPM throughout the autorotation to ensure RPM stays within limits.

When steady state autorotation is achieved, the pilot has the option of stretching his glide to a


6-10 AUTOROTATION

distant landing zone (LZ) or increasing his loiter time in the air, provided sufficient altitude

exists.

Rate of Descent and Glide Distance in Autorotation. Rather than using power required/power

available charts for autorotation, many NATOPS manuals contain charts specifically for

autorotation (Figure 6-9). Helicopter airspeed is probably the most significant factor that affects

rate of descent in autorotation. The rate of descent is high at very low airspeeds, decreases to a

minimum at some intermediate speed, and increases again at faster speeds. Minimum rate of

descent occurs at the bucket airspeed because this is where the minimum power is required to

remain airborne. If there is an available field immediately in front of you, you may use this

speed for extra time aloft to ensure crew readiness for landing or make a prudent radio

transmission, but there are other factors to consider as the helicopter approaches the ground.

If the engine failed and there was not a suitable landing site immediately in front of you, but

there was one further away, consider fling at maximum glide range airspeed. Maximum glide

distance occurs where the ratio of power required to airspeed is a minimum so that the aircraft

will fly the furthest horizontally with the smallest descent rate. A line drawn from the point of

origin tangent to the total drag curve illustrates the airspeed for maximum glide distance, much

the same as for powered flight. There are tradeoffs, and in this case, higher speed and distance

over the ground reduces time aloft.

The airspeeds for minimum rate of descent and maximum glide distance vary by helicopter type.

Individual operator’s manuals cover this information.

As the ground approaches the range of safe airspeed/rotor RPM combinations narrow and precise

management of kinetic energy is necessary. The new goal is to reduce kinetic energy along the

flight path to zero at the same time contact is made with the ground. Trade off the stored kinetic

energy in rotor RPM for thrust to maintain power requirements for flight before the blades reach

a stalled condition.

Deceleration and Touchdown. From either of the two extreme airspeed range examples

previously discussed (max glide/min rate of descent), we will assume a suitable LZ is now easily

within range. If we were at max glide at a high forward speed and associated high rate of

descent, it is only logical we slow down (low rate of descent at ground contact = less pain). How

slow? Minimum rate of descent sounds logical. But, even at this airspeed, the helicopter's

landing gear cannot absorb the amount of energy the helicopter is carrying at ground contact.

Therefore, it may be advantageous to carry five to ten knots extra airspeed over minimum rate of

descent airspeed at flare altitude - banking on another tradeoff – extra forward airspeed for high

rotor RPM. The fact is, maximum range airspeed is associated with greater rate of steady-state

descent, but the greater flare through minimum rate of descent airspeed to maximum decelerative

attitude provides a greater decelerative force both horizontally and vertically than a flare started

from min rate of descent speed. The challenge, therefore, is timing the deceleration

approximately from max range airspeed, which is more difficult than from the standard auto at

min rate of descent.


AUTOROTATION 6-11

Figure 6-9 Autorotational Rate of Descent Compared to Airspeed

The Last 100 Feet. It can be assumed that autorotation ends at 100 feet (depending on aircraft

and speed), and the landing procedure then begins. To execute a power-off landing for rotary-

wing aircraft, the aviator exchanges airspeed for lift by decelerating the aircraft during the last

100 feet. When executed correctly, deceleration is applied and timed so that rate of descent and

forward airspeed are minimized just before touchdown. At about 15 feet, this energy exchange

is essentially complete. The primary remaining control input is application of collective pitch to

cushion the touchdown. Because all helicopter types are slightly different, aviator experience in

that particular aircraft is the most useful tool for predicting the useful energy exchange available


6-12 AUTOROTATION

at 100 feet and the appropriate amount of deceleration and collective pitch needed to execute that

exchange safely and effectively to land the aircraft successfully.

Flare and Touchdown. Figure 6-10 shows the airflow and force vectors for a blade in an

autorotative deceleration. To make an autorotative landing, the aviator reduces airspeed and rate

of descent just before touchdown. The aviator can partially accomplish both actions by applying

aft cyclic, which changes the attitude of the rotor disk in relation to the relative wind. A nose-up

cyclic flare tilts the rotor disk rearward which inclines the resultant thrust of the rotor system

to the rear, slowing forward speed. It also increases AOA on all blades by changing the

direction of airflow through the rotor system. The resulting increase in AOA creates more lift

along with the lift vector becoming more vertical, which decreases rate of descent. Moreover,

the downward shift in relative wind tilts the lift vector at blade element more forward, resulting

in a larger pro-autorotative force; this increases rotor RPM. The increase in RPM can be used

to cushion the landing but must be monitored to prevent overspeeding the rotor head.

Additionally, the flare exposes more of the fuselage to the airstream, thereby increasing

fuselage parasitic drag, further aiding in slowing the aircraft down. The flare should be

maintained in an effort to reach a point where forward speed is five to ten knots at close

proximity to the ground (10-15 feet). At this point, increasing collective increases thrust (trading

RPM for lift) and augments braking action, using up part of the stored rotational energy. Due to

the aft-tilted thrust vector and the addition of collective, the pilot must put in a little forward

cyclic to level the aircraft and use that last rotational energy by pulling collective to cushion the

landing. Since there is no torque from the engine, drivetrain drag may cause the fuselage to

“follow” the rotor system when collective is pulled, causing the nose to yaw to the left and

requiring some right rudder, the opposite of powered flight. The key is to maintain heading

control throughout the autorotation using the rudder pedals as necessary.

Figure 6-10 Blade Element and Thrust during Steady State Auto and Flare


AUTOROTATION 6-13

If one chooses to arrive at flare altitude at less than minimum rate of descent airspeed, there is

little or no forward speed to trade off for this advantageous increase in rotor RPM and braking

action. Forward speed is already low, and if too much flare is combined with an improperly

timed flare (too high), forward speed may reduce to zero at a high altitude. This condition is

known as becoming “vertical,” and since the rotor system already has little stored energy, there

will not be enough thrust available with collective increase to slow rate of descent at touchdown

to a non-destructive level.

Arriving at the altitude for maximum decelerative effectiveness just prior to collective pull is

extremely useful because it involves trading kinetic energy from the descent for kinetic energy in

the rotor. If the attitude was achieved and held for an extended period of time, the reduction in

velocity would cause a greatly increased descent rate, which would increase the size of the stall

region on the blades and cause a loss of RPM.

Factors affected by the flare are: RPM, Thrust vector direction and magnitude, and Fuselage

attitude.

607. WINDMILL BRAKE STATE

Windmill Brake State. If the rotor somehow entered a descent at a rate in excess of

approximately 180% of induced velocity, too much potential energy would be diverted to

powering the rotor. Excessive rotor speed would create a very dangerous condition. In the

windmill brake state virtually all flow is “up” relative to the rotor, and energy may be extracted

from the system. This is the condition in which windmills extract energy from the passing air,

but it is not a normal operating state for any helicopter. The one time that a rotorcraft is put into

a descending flow state that approaches windmill brake is during the cyclic flare of an

autorotation. During that transient phase, kinetic energy is increased in the rotor by increasing

up-flow in order to make it available for use in the landing. However, RPM must be

monitored and controlled to prevent excessive buildup and overspeeding of the rotor head.

608. HEIGHT-VELOCITY DIAGRAM

No matter how well the pilot can execute an autorotation, there remain some combinations of

initial altitudes and airspeeds from which a safe autorotational landing will be extremely difficult

to perform. In fact, at some combinations of altitude and forward speed, it is almost impossible

to demonstrate safe autorotative landings at a vertical touchdown speed within the design limits

of the landing gear. The boundaries of these combinations define the height-velocity diagram or

"The Deadman's Curve." (Figure 6-11)

The purpose of an H–V diagram is to identify the portions of the flight envelope from which a

safe landing can be made in the event of a sudden engine failure. The H–V diagram also

generally depicts two areas to be avoided: The low-airspeed/high-hover altitude (low flight

altitude) region and the high-airspeed/low-altitude region. These are named with respect to

takeoff from the IGE Hover. At a hover, 200 – 300’ is considered a high altitude. Above

60 KIAS, flight below 20’ is considered low altitude.


6-14 AUTOROTATION

There are H-V diagrams for each type of helicopter. They are found in their respective NATOPS

manuals. Helicopter pilots should be familiar with these diagrams.

Taking a closer look at the H-V diagram, we see several definite points define the curve, the first

being the low hover height. Up to this height, a pilot can handle a power failure by coming

straight down, using collective increase to cushion the landing. Above that altitude in

combination with low speed, the rotor blades will slow down and stall if collective setting

remains constant, or the helicopter will impact the ground too hard if collective is lowered.

Enough altitude does not exist to acquire enough forward airspeed by the time flare altitude is

reached to successfully execute a flare. This height is a function of the power required to hover,

rotor inertia, blade area and stall characteristics, and the capability of the landing gear to absorb

the landing forces without sustaining damage.

The unsafe hover area runs from the low hover height to the high hover height. Above this

altitude, there is enough altitude to make a diving transition into forward flight autorotation and

execute a normal flare.

Beyond the knee of the curve, a power failure is survivable at any altitude above the high-

airspeed/low-altitude region. The three problems associated with the high-airspeed/low-altitude

region are pilot reaction time, lack of time and altitude for the induced flow to reverse before

ground impact, and possibility of tail rotor stinger strike in response to cyclic flare to trade

altitude for airspeed.

Figure 6-11 Generic Height Velocity Diagram

Skilled test pilots, who try to make their reactions simulate those of the average reaction time of

a pilot, establish H-V diagrams. This is done by specifying a definite delay time following the


AUTOROTATION 6-15

engine failure before initiating control input. The military assumes their pilots may be distracted

during an engine failure due to focused attention to assigned missions, allowing a two-second

delay before response during any flight condition.

A few regions of operation are apparent on the H-V diagram.

1. Low-speed region. The largest region noticed on the diagram occurs where potential

energy is not sufficient to offset low aircraft kinetic energy state for transition to an autorotative

glide path. In other words, not enough altitude is available to establish a steady-state glide and

minimum flare airspeed. When the engine fails in this sector a rapid rate of descent will occur,

but with little or no forward airspeed a flare will not be capable of arresting the descent prior to

landing. Application of collective will cause the RPM to decay excessively resulting in a hard

landing (limited rotor rotational kinetic energy available).

2. Low-altitude/high-speed region. At low altitude and high speed, a quick cyclic flare can

transfer kinetic energy to the rotor, provided time is sufficient to initiate the maneuver before

ground impact. In the low-altitude/high-speed region, velocity is too great for a safe taxiing

auto, but altitude is too low for flare initiation. By the time the pilot reacts (using typical

reaction time) with a flare or zoom climb, the tail sinks enough to impact the ground.

3. High hover height. At altitudes above the low airspeed avoid region, the pilot can enter

autorotation by making a diving transition to forward flight, reaching the desired autorotation

airspeed and then executing a normal flare.

4. Low hover height. Below the low airspeed avoid region the helicopter can simply be

landed straight down with no forward airspeed and cushioned with collective and/or landing

gear. Rotor stored kinetic energy is traded in the cushion.

The size of the avoid region is affected by several variables. Pilot response time varies, but

charts are drawn on the basis of average pilot response times. Airspeed affects the ability to

establish an autorotation at an acceptable rate of descent. Rotor inertia determines how quickly

the rotor loses speed. A low-inertia system would be more likely to lose valuable RPM before

establishment of autorotation, so it would have a larger avoid area than a helicopter with high

rotor inertia. Increased gross weight increases power required for flight, which in turn increases

rate of descent and thus increases the size of the avoid area. DA decreases rotor efficiency and

increases power required, so it has the same effect as increased gross weight on size of the avoid

area.

The variables that directly affect the size of the height-velocity diagram avoid area are:

1. Rotor Inertia. High inertia reduces shaded region because RPM does not decay as fast as

a low-inertia system. High inertia moves the “knee” left.

2. Gross Weight. Power requirements increase. High gross weight moves the “knee” right.

3. DA. Same effect as gross weight. High DA moves the “knee” right.


6-16 AUTOROTATION

H-V Guidelines in the TH-57. During normal takeoff, airspeed should be 40 KIAS by 20’ AGL

and 65 KIAS by 50’ AGL to minimize the risk near the H-V diagram avoid areas, transitioning

through the caution area into the green area as quickly as possible. In Chapter Four of NATOPS,

protracted operations in the AVOID and CAUTION areas of the height-velocity diagram are

prohibited. Realize that mission operations such as rescue hoisting, externals, etc., all require

“high” altitude hovers. One just needs to realize that if an engine failure occurs during those

operations, options may be limited!

PERFORMANCE 7-1

CHAPTER SEVEN

PERFORMANCE

700. INTRODUCTION

This chapter correlates helicopter aerodynamics principles with engine and aircraft performance.









702. FACTORS AFFECTING PERFORMANCE

Moisture (Humidity): Humidity alone is usually not considered an important factor in

calculating density altitude and helicopter performance; however, it does contribute. There are

no rules of thumb used to compute the effects of humidity on density altitude, but some

manufacturers include charts with 80 percent relative humidity columns as additional

information. There appears to be an approximately 3–4 percent reduction in performance

compared to dry air at the same altitude and temperature, so expect a decrease in hovering and

takeoff performance in high humidity conditions. Although 3–4 percent seems insignificant, it

can be the cause of a mishap when already operating at the limits of the helicopter.

Weight: Weight is one of the most important factors because the pilot can control it. Most

performance charts include weight as one of the variables. By reducing the weight of the

helicopter, a pilot may be able to take off or land safely at a location that otherwise would be

impossible. However, if ever in doubt about whether a takeoff or landing can be performed

safely, delay your takeoff until more favorable density altitude conditions exist. If airborne, try

to land at a location that has more favorable conditions, or one where a landing can be made that

does not require a hover.

In addition, at higher gross weights, the increased power required to hover produces more torque,

which means more anti-torque thrust is required. In some helicopters during high altitude

operations, the maximum anti-torque produced by the tail rotor during a hover may not be

sufficient to overcome torque even if the gross weight is within limits.

CHAPTER SEVEN HELICOPTER AERODYNAMICS WORKBOOK

7-2 PERFORMANCE

Winds: Wind direction and velocity also affect hovering, takeoff, and climb performance.

Translational lift occurs any time there is relative airflow over the rotor disk. This occurs

whether the relative airflow is caused by helicopter movement or by the wind. Assuming a

headwind, as wind speed increases, translational lift increases, resulting in less power required to

hover.

The wind direction is also an important consideration. Headwinds are the most desirable as they

contribute to the greatest increase in performance. Strong crosswinds and tailwinds may require

the use of more tail rotor thrust to maintain directional control. This increased tail rotor thrust

absorbs power from the engine, which means there is less power available to the main rotor for

the production of lift. Some helicopters even have a critical wind azimuth or maximum safe

relative wind chart. Operating the helicopter beyond these limits could cause loss of tail rotor

effectiveness.

Takeoff and climb performance is greatly affected by wind. When taking off into a headwind,

effective translational lift is achieved earlier, resulting in more lift and a steeper climb angle.

When taking off with a tailwind, more distance is required to accelerate through translation lift.

703. GENERAL

Helicopter aircraft and engine performance require an understanding of the power required

curves, power available, and the relationship between them.

Power is required to overcome the drag produced by the rotors and the fuselage. The power

available to meet this power requirement is produced by a turboshaft engine. Turbojet aircraft

produce thrust directly from their engines and do not turn a propeller or rotor. As such, jet

aircraft performance charts only slightly resemble helicopter drag curves. For each pound of

drag generated by the aircraft at a specific airspeed, a pound of thrust must be generated by the

jet in order to maintain level flight. The amount of thrust produced by a jet engine is directly

proportional to fuel flow and therefore endurance and range performance may be determined

from an aircraft total drag curve.

The differences between the two types of performance curves can be attributed to the different

contributions of profile and induced drag in the helicopter. The helicopter rotor also produces

thrust, but the production of thrust is not directly related to the fuel flow for turbo-shaft engines.

Turbo-shaft fuel flow is more closely related to how much power is being produced by the

engine. Accordingly, a total drag curve cannot be used in the same way as a performance chart

for helicopters.

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER SEVEN

PERFORMANCE 7-3

Figure 7-1 Total Drag Curve

Instead, “power required curves” (or more specifically, “fuel flow curves”) are presented in

operator’s manuals for use in mission planning. Power is simply the rate of doing work. Most

turbine helicopters are equipped with a gauge for measuring torque which may be viewed by the

aviator in the cockpit. Since power equals torque times RPM, if the RPM remains constant, the

torque is a direct representation of current engine power output. Further, a fuel flow scale is

usually provided opposite the torque scale of a cruise chart, thereby enabling the aviator to

convert torque directly to fuel flow.


7-4 PERFORMANCE

Figure 7-2 Power Required Versus Airspeed Curve

It is important to note that the lowest point on the power required curve is the point of minimum

power required (best lift to drag ratio) and not necessarily the point of minimum drag (as is the

lowest point on the total drag curve). The point of minimum power results in the lowest fuel

flow and is therefore the airspeed for maximum endurance. The airspeed for minimum power

is slower than the airspeed for minimum drag because a decrease in velocity to the minimum

power airspeed decreases the power required, even though flying at any airspeed below

minimum drag actually increases drag. However, because the bottom of the drag curve is nearly

flat, the slight increase in drag is more than offset by the decrease in velocity, which slows the

work rate and therefore results in an overall reduction in power required.

704. POWER REQUIRED

For a helicopter to remain in steady, level flight, forces and moments must balance. These forces

exist in the vertical plane, horizontal plane, and about the CG in the form of pitching moments.


PERFORMANCE 7-5

Figure 7-3 Aerodynamic Forces Affecting Power Required

Figure 7-4 Power Required Curves versus Airspeed

In a hover, two types of power are necessary - induced and profile power. Induced power is

power associated with the production of rotor thrust. This value is at its highest during a hover

(60 – 85 percent of total main rotor power) and decreases rapidly as the helicopter accelerates

into forward flight. During forward flight, the increase in mass flow of air introduced to the

rotor system reduces the amount of work the rotors must produce to maintain a constant thrust,

therefore, induced power required continues to go down with increasing airspeed.

Profile power, which can be thought of as "main rotor turning power," accounts for 15 – 40

percent of main rotor power in a hover and is used to overcome friction drag on the blades. It

increases slightly with increasing airspeed.


7-6 PERFORMANCE

In forward flight, parasite power joins forces with induced and profile power to overcome the

parasite drag generated by all the aircraft components, excluding the rotor blades. Parasite

power can be thought of as the power required to move the aircraft through the air. This power

requirement increases in proportion to forward airspeed cubed. Obviously, this is

inconsequential at low speed, but is significant at high speed and is an important consideration

for helicopter designers to minimize drag.

In addition to the drag curves which are the basis for the power required curves, there is a fourth

power requirement, labeled “miscellaneous,” which is taken into account when power required

curves are developed for specific rotorcraft. This is the power required to run the tail rotor and

accessories such as generators, hydraulics, etc. Accessory power requirements remain relatively

constant independent of airspeed, while tail rotor power required tends to decrease with

increasing airspeed. Depending on the charts used, this additional power requirement is

sometimes combined with the profile power requirement, creating a “total rotor profile power”

required to maintain a given rotor RPM, taking into account the rotor profile drag as well as the

tail rotor and accessory requirements.

The smaller horizontal force, H-force, is produced by the unbalanced profile and induced drag

(or in-plane drag in some books) of the main rotor blades. Tilting the rotor disk forward from a

fraction of a degree at low speed to about 10° at max speed compensates for this.

Different flight regimes are performed more efficiently at different forward speeds. The bowl-

shape of the power required curve graphically illustrates the reason why. Optimum speeds

determined by this curve are maximum loiter time (endurance), minimum rate of descent in

autorotation, best rate of climb and maximum range airspeed.

Best rate of climb airspeed is formed at the point where the difference is a maximum between

power required and power available. The bottom of the curve is called the bucket airspeed.

Since the goal of achieving maximum loiter time is making the available fuel last as long as

possible, and since fuel flow is proportional to engine power, maximum loiter time should also

be at this point.

Near this speed, minimum rate of descent in an autorotation is also found, since the power

required to keep the aircraft airborne is at a minimum.


PERFORMANCE 7-7

Figure 7-5 Optimum Airspeeds

The point on the power required curve corresponding to the point of minimum drag versus

airspeed on the drag curve is at an airspeed greater than that for minimum power (bucket

airspeed). This is the airspeed for maximum range and is where the ratio of fuel flow to

velocity is at a minimum value. This point is shown in at the point of tangency of the power

required curve and a straight line drawn from the origin, providing the best power or fuel flow to

airspeed (thus drag) ratio. Maximum range speed is found on the fuel flow curve by drawing a

line tangent to the curve from the origin. This ratio of speed to fuel flow shows the distance one

can travel on a pound of fuel on a no-wind day. If there is a headwind, the line should be

originated at the headwind value, which derives a higher speed and lower range. For a tailwind,

the optimum airspeed decreases, but the range increases significantly. On generic charts the

speed for maximum range and autorotation maximum glide distance sometimes appear to be the

same. Best airspeed for maximum glide in an autorotation is also affected by headwinds and

tailwinds just like maximum range airspeed. However, when using aircraft specific charts, that

is not usually the case and the two speeds are different.


7-8 PERFORMANCE

Figure 7-6 Maximum Range Airspeed Adjustment for Winds

705. POWER AVAILABLE

Air density (DA) is the environmental factor which most significantly affects power available.

Less dense air requires that the engine works harder to produce the same amount of mass flow.

Power available is directly affected by density to such a degree that power available at a given

DA can be calculated by simply multiplying power available at standard sea level by the density

ratio in the ambient conditions.

Poweravail = Powersea level x density ratio ()

This is generally true in regions of relatively normal temperature variation. However, in

locations of extremely wide temperature variations such as the desert environment, the

temperature can have an extra degrading effect on engine power available.

Other factors limiting power available. Operating conditions that affect fuel flow or airflow

directly affect the ability of the engine to generate power. Some of those factors follow:

1. Fuel Flow Limitation (cold) - As temperature decreases the density of air increases so the

fuel flow must increase in order to maintain the stoichiometric fuel/air ratio for complete

combustion. However, the amount of fuel flow through the fuel nozzles has a limit; therefore at

cold temperatures the fuel/air ratio will not be optimum, and incomplete (lean) combustion will

occur, resulting in less power available.


PERFORMANCE 7-9

2. Turbine Temperature Limitation (hot) - The materials used to build turbines have definite

stress and temperature limits. To avoid unacceptable creep or component failure, turbine

temperature must be limited. Depending on aircraft manufacturer this can be called exhaust gas

temperature (EGT), turbine outlet temperature (TOT) or turbine inlet temperature (TIT).

3. Ng-Gas Generator Limitation (hot) - As the OAT increases the density of air decreases,

therefore the gas turbine has to rotate faster in order to deliver the same mass flow rate. This

increased rotational speed required at higher temperatures can approach limits that have been

established to counter centrifugal loads on the gas turbine blades.

4. Age of the engine - Compressor blades erode with time, and their degradation results in

loss of blade area that will degrade engine performance.

5. Component rating degradation - Transmission components have material limitations, so

their torque capacity must be considered.

6. Humidity/Moisture Effect - Increases in humidity/moisture have a counteracting effect in

that the associated decrease in air density is detrimental while the reduced combustor inlet temp

(T3) is beneficial. Hence humidity/moisture has a negligible effect on gas turbine engines.

7. Torque limits - drive train limits, including drive shaft and transmission.

8. Airspeed effects (ram air) - Airspeed increases the flow rate into the engine, but at the

speed at which rotorcraft operate this effect is negligible.

706. EFFECT OF TAIL ROTOR ON POWER AVAILABLE

Since the engine drives both the main rotor and the tail rotor, the tail rotor does not affect the

power available that the engine produces; rather, it requires that the power be shared between the

two rotor systems, reducing power available for the main rotor. The tail rotor uses 5-15% of the

total power available, therefore leaving only 85-95% for the main rotor. Although other

frictional losses of the drive train may be significant, the tail rotor robs the greatest amount of

power from the main rotor. The tail rotor makes its greatest demands on the engine power

available when the greatest requirements are on the main rotor. For example, in the climb,

termination of a steep approach, when power required to perform a maneuver is equal to or

exceeding the power available, and/or when rotor RPM is drooping, the main rotor system is

creating the greatest amount of torque, therefore, the anti-torque requirements from the tail rotor

are greatest. However, since the main rotor and tail rotor are driven by a common system, when

main rotor RPM droops, tail rotor demand is highest and it is most sensitive to its own decreased

RPM.

For each of these maneuvers, the time to insure adequate power is available to maintain tail rotor

and main rotor authority is prior to the maneuver. This highlights the critical importance of in-

depth performance planning prior to flight, as well as careful re-evaluation during flight should

the mission require any alteration to the planned flight. When the margin for error is minimal,

unnecessary maneuvering should be kept to a minimum, and increased vigilance is required to be


7-10 PERFORMANCE

best prepared for any unanticipated situation. As with any other aspect of aviation, expect the

unexpected, and then vortex ring state (VRS), PR Exceeds PA, or even an engine failure will not

catch you unprepared.

707. POWER REQUIRED EXCEEDS POWER AVAILABLE

When power required for a maneuver exceeds power available (PR>PA) under the ambient

conditions, an un-commanded descent or deceleration will result. Aggravating factors include:

high G-loading, high gross weight, high DA, rapid maneuvering (quick stops), slow spool up

time, loss of wind effect, loss of wind direction, and loss of ground effect (transiting from the

deck of a ship).

PR>PA Indications. Power required exceeding power available becomes dangerous when

operating in close proximity to obstructions where the pilot may not have maneuvering airspace

to recover prior to impacting the obstacle. Along with the un-commanded descent or

deceleration, rotor droop and associated loss of tail rotor authority (LTA) may result.

In addition to proper performance planning and situational awareness of the above aggravating

factors, the pilot should avoid excessive maneuvering, high descent rates, and downwind

takeoffs and landings, especially in environmental conditions where power available may be

marginal.

Power required exceeding power available is differentiated from vortex ring state (VRS) by un-

commanded descent being associated with max allowable torque and/or rotor droop and possible

LTA. VRS is not normally associated with either rotor droop or LTA.

Induced power. Induced power requirements change as forward airspeed increases. The

requirement for a mass flow of air still exists, but forward velocity increases the mass flow rate

so the rotor does not need to apply as much work on the air, thus less power is required. At

speeds beyond that at which the tip vortices are outrun (speed for translational lift) the rotor disk

acts in a manner that is similar to a conventional wing.


PERFORMANCE 7-11

Figure 7-7 Induced Power Required

Parasite power. Just as high induced power requirements can cause a power required exceeding

power available situation at a hover or slow airspeeds, parasite power requirements can cause the

situation at higher speeds.

Figure 7-8 Parasite Power Required

The point where power required exceeds engine power available at high speed is commonly

referred to as VH or the maximum speed in level flight at maximum power without an un-

commanded descent developing. VH should not be confused with VNE (Velocity never exceed)

which is a structural limitation. VNE of an aircraft is the V speed which refers to the velocity that

should never be exceeded due to risk of structural failure, due to calculated factors such as wing,

tail, or airframe deformation or due to aero elastic 'flutter' (unstable airframe or control

oscillation). VNE is specified as a red line on many airspeed indicators. This speed is specific to

the aircraft model, and represents the edge of its performance envelope in terms of speed.

Excess Power. Because power required exceeding power available is often associated with


7-12 PERFORMANCE

lower airspeeds, the induced power requirement may become critical. As airspeed decreases

during the approach or maneuvering, induced power required increases, so deceleration with a

constant collective setting moves the helicopter into a regime where excess power (difference

between available and required) steadily decreases. Each aircraft is inherently different, and

experiences a different amount of translational lift at different airspeeds. The determining factor

in the amount of "extra lift" available is determined by the slope of the induced power required

curve. Typically, translational lift is experienced during transition from hover into forward flight

at airspeeds between 13 - 24 knots, depending on disk size, blade area, and RPM.

Figure 7-9 Decrease in Excess Power as Airspeed Decreases

It is a fact that when a helicopter transitions from forward flight to a hover, it experiences

decreased performance because of increased induced power requirements that stem from the tip

vortices that are generated in a hover. As airspeed decreases to near 13 - 24 knots the entire rotor

begins to experience recirculation of vortices, and vortices impact the fuselage and tail. Power

required increases and if power available is marginal, or the aircraft is not in ground effect,

conditions are ripe for an un-commanded descent, rotor droop, and/or LTA.

708. HOVER PERFORMANCE

Helicopter performance revolves around whether or not the helicopter can be hovered. More

power is required during the hover than in any other flight regime. Obstructions aside, if a hover

can be maintained, a takeoff can be made, especially with the additional benefit of translational

lift. Hover charts are provided for in ground effect (IGE) hover and out of ground effect (OGE)

hover under various conditions of gross weight, altitude, temperature, and power. The IGE

hover ceiling is usually higher than the OGE hover ceiling because of the added lift benefit

produced by ground effect. A pilot should always plan an OGE hover when landing in an area

that is uncertain or unverified.

As density altitude increases, more power is required to hover. At some point, the power


PERFORMANCE 7-13

required is equal to the power available. This establishes the hovering ceiling under the existing

conditions. Any adjustment to the gross weight by varying fuel, payload, or both, affects the

hovering ceiling. The heavier the gross weight, the lower the hovering ceiling. As gross weight

is decreased, the hover ceiling increases.

Sample Hover Problem 1

You are to fly to a remote location for training. Using Figure 7-10 can you safely hover in

ground effect at your departure point with the following conditions?

Pressure Altitude..............................8,000 feet

Temperature..........................................+15 °C

Takeoff Gross Weight.........................1,250 lb.

RPM...............................................104 percent

First enter the chart at 8,000 feet pressure altitude (point A), then move right until reaching a

point midway between the +10 °C and +20 °C lines (point B). From that point, proceed down to

find the maximum gross weight where a 2 foot hover can be achieved. In this case, it is

approximately 1,280 pounds (point C). Since the gross weight of your helicopter is less than

this, you can safely hover with these conditions.


7-14 PERFORMANCE

Figure 7-10 Sample hover problem 1


Once you reach the remote location in the previous problem, you will need to hover OGE for

training. The pressure altitude at the remote site is 9,000 feet, and you will use 50 pounds of fuel

getting there. (The new gross weight is now 1,200 pounds.) The temperature will remain at

+15 °C. Can you accomplish the mission?

Enter the chart at 9,000 feet (point A) and proceed to point B (+15 °C). From there, determine

that the maximum gross weight to hover OGE is approximately 1,130 pounds (point C). Since

your gross weight is higher than this value, you will not be able to hover in these conditions. To

accomplish the mission, you will need to remove approximately 70 pounds before you begin the

flight.


PERFORMANCE 7-15


These two sample problems emphasize the importance of determining the gross weight and

hover ceiling throughout the entire flight operation. Being able to hover at the takeoff location

with a specific gross weight does not ensure the same performance at the landing point. If the

destination point is at a higher density altitude because of higher elevation, temperature, and/or

relative humidity, more power is required to hover there. You should be able to predict whether

hovering power will be available at the destination by knowing the temperature and wind

conditions, using the performance charts in the helicopter flight manual, and making certain

power checks during hover and in flight prior to commencing the approach and landing.

For helicopters with dual engines, performance charts provide torque amounts for both engines.


7-16 PERFORMANCE


Determine what torque is required to hover. Use the following conditions:

Pressure Altitude ............................................9,500 feet

Outside Air Temperature.........................................0 °C

Gross Weight.....................................................4,250 lb.

Desired Skid Height ..............................................5 feet

First, enter the chart at 9,500 feet pressure altitude, then move right to outside air temperature,

0 °C. From that point, move down to 4,250 pounds gross weight and then move left to 5-foot

skid height. Drop down to read 66 percent torque required to hover.



PERFORMANCE 7-17

709. CLIMB PERFORMANCE

Most of the factors affecting hover and takeoff performance also affect climb performance. In

addition, turbulent air, pilot techniques, and overall condition of the helicopter can cause climb

performance to vary.

A helicopter flown at the best rate-of-climb speed (VY) obtains the greatest gain in altitude over a

given period of time. This speed is normally used during the climb after all obstacles have been

cleared and is usually maintained until reaching cruise altitude. Rate of climb must not be

confused with angle of climb. Angle of climb is a function of altitude gained over a given

distance. The VY results in the highest climb rate, but not the steepest climb angle, and may not

be sufficient to clear obstructions. The best angle of climb speed (VX) depends upon the power

available. If there is a surplus of power available, the helicopter can climb vertically, so VX is

zero.

Wind direction and speed have an effect on climb performance, but it is often misunderstood.

Airspeed is the speed at which the helicopter is moving through the atmosphere and is unaffected

by wind. Atmospheric wind affects only the groundspeed, or speed at which the helicopter is

moving over the Earth’s surface. Thus, the only climb performance affected by atmospheric

wind is the angle of climb and not the rate of climb.

When planning for climb performance, it is first important to plan for torque settings at level

flight. Climb performance charts show the change in torque, above or below torque, required for

level flight under the same gross weight and atmospheric conditions to obtain a given rate of

climb or descent.

Sample Cruise or Level Flight Problem

Determine torque setting for cruise or level flight using the following conditions:

Pressure Altitude............................................. 8,000 feet

Outside Air Temperature..................................... +15 °C

Indicated Airspeed...............................................80 knots

Maximum Gross Weight .....................................5,000 lb.

With this chart, first confirm that it is for a pressure altitude of 8,000 feet with an OAT of 15°.

Begin on the left side at 80 knots indicated airspeed (point A) and move right to maximum gross

weight of 5,000 lb. (point B). From that point, proceed down to the torque reading for level

flight, which is 74 percent torque (point C). This torque setting is used in the next problem to

add or subtract cruise/descent torque percentage from cruise flight.


7-18 PERFORMANCE

Figure 7-13 Sample cruise problem

Sample Climb Problem

Determine climb/descent torque percentage using the following conditions:

Rate of Climb or Descent .....................................500 fpm

Maximum Gross Weight......................................5, 000 lb.

With this chart, first locate a 500-fpm rate of climb or descent (point A), and then move to the

right to a maximum gross weight of 5,000 lb. (point B). From that point, proceed down to the

torque percentage, which is 15 percent torque (point C). For climb or descent, 15 percent torque

should be added/subtracted from the 74 percent torque needed for level flight. For example, if

the numbers were to be used for a climb torque, the pilot would adjust torque settings to

89 percent for optimal climb performance.


PERFORMANCE 7-19

Figure 7-14 Sample climb problem

710. REVIEW OF OPTIMUM AIRSPEEDS

Optimum Airspeeds. Choosing appropriate level-flight airspeed is an important part of

obtaining the most appropriate performance for a mission. Airspeeds for maximum speed,

maximum range, and maximum endurance are distinct, and vary with aircraft loading and

environment. In a given set of conditions, one performance parameter may be more crucial than

others, so flight at the airspeed that would maximize that potential makes sense. For example, if

holding while awaiting deck landing space is important, the pilot should fly at the airspeed that

gives the best endurance. In a long over-water mission, best range may be appropriate.

Maximum speed is required in a time critical situation. Fortunately, the required airspeed for

any of these situations is easily found on a power required versus airspeed chart.


7-20 PERFORMANCE

Figure 7-15 Optimum Airspeeds

Speeds for endurance and range. As a previous lesson pointed out, power required varies with

weight, altitude, and airspeed. The chart section of a NATOPS manual contains power required

curves that are divided up typically by PA (pressure altitude). They depict power requirements

at a variety of aircraft weights in a format similar to that used here. The shape of this curve has

been likened, by some, to be just like a collective position curve; the collective is highest in a no-

wind hover, decreases with increasing forward airspeed to the "bucket," and increases again as

airspeed approaches VNE.

The power required curve also depicts fuel flow required at various airspeeds because power has

a direct relationship to the amount of fuel introduced into a gas turbine. In a no-wind hover,

power required is the highest, so fuel flow is also the highest. Power and fuel flow decrease as

forward velocity is increased toward the "bucket,” and then increase again as airspeed

approaches VNE. Thus, a relationship between fuel flow and forward speed can be visualized.

Minimum fuel flow occurs at the bucket airspeed, so the bucket airspeed is identified as the point

of maximum endurance.

The airspeed for maximum range is determined by drawing a tangent line from the origin to the

fuel flow/power required curve. The slope, or the change in the vertical direction with respect to

the change in the horizontal direction, is fuel flow/airspeed. Units of the slope are lb. /hr divided

by NM/hr. When the hours are cancelled in the slope, the units of slope are pounds of fuel used

per nautical mile traveled. Minimizing the slope translates to finding the point at which the least

fuel is burned for each nautical mile of travel (pound/NM). The least fuel per nautical mile is the

same as the most distance covered for the least amount of fuel used, or maximum range.


PERFORMANCE 7-21

Figure 7-16 Power Required Chart (CH-46E)

So how do winds factor into this situation? Winds do not affect endurance airspeed because

distance over the ground is not important in endurance calculations. Endurance solely deals with

time aloft at minimum power, minimum amount of fuel burned, and airspeed felt at the rotor is

the same whether wind is present or not.

Maximum range, however, is affected by winds because it involves maximizing movement over

the ground with minimal fuel flow. Speed over the ground is faster with a tailwind and slower

with a headwind, so the origin (zero point) of the airspeed axis must be shifted by the amount of

the headwind. Thought of another way, with a headwind the aircraft does not go as far and with

a tailwind it goes further.

Repeated use of the above system of determining wind correction has given results at typical

maximum range airspeeds that are predictable even without chart work. A good rule of thumb,

based upon consistently close approximations, is to add 1/4 of the headwind component on to no-

wind maximum range airspeed and to subtract 1/6 of the tailwind component. It should be noted

that these percentages will change with gross weight, ambient air conditions, and aircraft T/M/S.

The aviator is recommended to thoroughly sift through his own power and airspeed charts in

order to validate these trends. The difference between the two adjustments has to do with the

shape of the curve and the effect of shifting the origin and the resulting point of intersection of

the tangent line.

Maximum rate of climb and minimum rate of descent. The airspeed to fly for maximum

endurance (the bucket airspeed) is also suitable for maximizing performance in two other

regimes: maximum rate of climb and minimum rate of descent with power off. Maximum rate

of climb occurs at the bucket because minimum power required subtracted from a fairly constant

power available yields the largest amount of excess power available. If airspeed is maintained

constant, excess power can be used to climb or maneuver. Likewise, in a descent, the point of

minimum power required for flight is the point at which power deficit, which relates directly to

sink rate, would occur in a power-off situation. Note that max endurance and max rate of climb


7-22 PERFORMANCE

are found at the bucket of a power chart, while minimum rate of descent is found at the bucket of

an autorotation chart.

Figure 7-17 Fuel Flow vs. TAS

Figure 7-18 Excess Power

How does fuel consumption change with altitude? The specific fuel consumption of the gas

turbine varies with two primary operating parameters: temperature and power output. The

specific fuel consumption is defined as nautical miles per pound of fuel.


PERFORMANCE 7-23

As density is decreased, both the fuel flow and SHP decrease proportionally, so it can be said

that density variations do not by themselves influence specific fuel consumption (disregarding

profile and parasitic drag). However, as altitude is increased, temperature normally decreases.

Because the turbine may deliver a given thrust output with less fuel at a lower inlet temperature,

the specific fuel consumption normally improves (decreases) with altitude. If the atmosphere

can be considered to be standard, the specific fuel consumption decreases to the tropopause and

then remains constant until the efficiency of the compressor begins to break down at sufficiently

high altitudes. The standard atmosphere has a temperature decrease up to the tropopause

(approx. 36,000 feet).

Figure 7-19 Max Range Altitude vs. Gross Weight

Specific fuel consumption also varies quite noticeably with power output. The gas turbine is

designed so that it operates most efficiently at high power outputs. This means that the specific

fuel consumption is lowest at higher power settings, and that 100% Ng is the optimum speed for

greatest efficiency. Note that total fuel consumption does not go down at high power settings;

specific fuel consumption, or pound of fuel per hour per SHP does. For a given amount of shaft

horsepower output, the least amount of fuel is burned (highest efficiency) at high power settings.

This situation poses an interesting problem for helicopters. Because helicopters use turboshaft

engines, the fuel efficiency of the aircraft is determined by the efficiency of both the engine and

the rotor system. An increase in DA requires more work by the rotor system for the same flight

profile. Engine efficiency gains at altitude are balanced by rotor efficiency losses. Actual fuel

efficiency obtained at altitude thus depends upon rotor system efficiency, installed aircraft

engine characteristics, work output requirement, and total fuel load. In general, at low gross

weights one gets better range at higher altitude while at high gross weights a better range is

achieved at sea level.


7-24 PERFORMANCE

As might be expected, fuel flow increases at higher gross weights. As aircraft gross weight

increases, power required increases and hence fuel flow increases. Also, the airspeed for

maximum range also increases due to the shift in the power required curve. Maximum range and

endurance airspeeds increase with increasing gross weight. This is true because increased gross

weights shift the power required curve up and to the right. This trend is universally true for

maximum endurance airspeeds, but varying shapes of the power required curve for some

helicopters make the trend of best range airspeed unpredictable. A survey of several fleet

helicopters reveals that maximum range airspeed shifts depend upon the aircraft and operating

environment.

For example, the AH-1W and MH-60S show decreased maximum range airspeed at higher gross

weights for the conditions given.

Figure 7-20 AH-1W - Max Range Airspeed vs. Gross Weight


PERFORMANCE 7-25

Figure 7-21 MH-60S – Max Range Airspeed vs. Gross Weight

Effect of rotor speed on range and endurance. In some situations, a helicopter may consume

less fuel at a rotor speed below 100%. This benefit only occurs when profile power is a major

contributor to power required, so it only applies to a certain extent. When the rotor speed gets

too slow the increase in AOA required to generate lift at a slower rotational speed generates

excessive drag forces.

In addition to possible drag increases, decreasing Nr for fuel efficiency can present other

problems. Decreased main rotor speed produces lower tail rotor speed, so loss of tail rotor

efficiency can increase power demands and, in the most extreme case, make LTA more likely.

Additionally, in the event of an engine failure, rotor RPM will decay to dangerous levels more

quickly.

Nonetheless, it is true that in some cases a decrease in Nr can yield a decreased fuel flow that will

increase range and endurance. Figure 7-22 shows that fuel flow increases below 100% Nr at

high gross weights because the rotor system is attempting to lift a heavier aircraft at slower than

optimal rotor speed. This feature is overwhelmingly obvious with the CH-53D but the trend is

the same for all aircraft.

Thus, fuel conservation benefits occur primarily at lower percentages of maximum gross weight.

Even at lower gross weights, use of this technique should be carefully considered for its

necessity, thoroughly planned, and not used routinely. One hundred percent Nr is established by


7-26 PERFORMANCE

Figure 7-22 RPM vs. Fuel Flow

Another use for Bucket Airspeed. Several fleet helicopter mishaps involving gearbox failures

have addressed the selection of an airspeed to fly in cases of impending catastrophic component

failure while flying over water. NATOPS manuals provide some guidance in this area, and it is

no coincidence that recommended speeds are generally in the vicinity of that recommended for

maximum endurance. The best airspeed would be that which demands the lowest power

requirements, thereby imposing the smallest load on the defective transmission/gearbox

components in an attempt to delay failure. The smallest load occurs at the bucket airspeed,

which is the same as the airspeed for maximum rate of climb, airspeed for minimum rate of

descent in a power-off situation, and airspeed for maximum endurance. In at least one case

(the CH-53E), an airspeed range is recommended for main gear box oil system failure, but the

text recommends specifically that airspeed be reduced to minimum power required for flight: the

bucket.

For a given gross weight and altitude, one airspeed, the bucket, provides the performance point

which maximizes potential climb rate for achieving communication/navigation reception,

minimizes fuel flow if required to loiter until assistance arrives and/or a positive fix is obtained,

provides for a minimum rate of descent in the event of a power loss and subsequent ditching, and

minimizes the mechanical loads on the failing components to the point of possibly delaying their

demise long enough to reach terra firma or some other suitable platform.


PERFORMANCE 7-27

The bucket is a very useful airspeed to keep in mind. Although the bucket shifts a little with

changes in gross weight (fuel burn-off) and altitude, an average speed target that works well for

the most common operating profiles is worthy of remembering. In fact, some NATOPS manuals

define a "best" airspeed that implies some average value over the normal gross weight/altitude

operating range.

Of course, nothing supersedes sound judgment and good headwork in adjusting to a particular

emergency situation. Selection of the most favorable airspeed can depend on desired closure

rates, maximum range requirements or other considerations once catastrophic failure is

imminent.

Excess Power. Power available (PA) is almost constant throughout all velocities, hence,

maximum excess power occurs at the "bucket" airspeed where power required (PR) is a

minimum. The airspeed for maximum excess power equates to airspeed for maximum rate of

climb, as previously discussed.

Figure 7-23 Excess Power

Best Angle of Climb. Excess power in a HOGE means that the best angle of climb is straight

up. If power is not sufficient for a vertical climb, then the best angle of climb occurs at an

airspeed that involves maximum vertical velocity per unit of horizontal velocity. That airspeed

is obtained by drawing a tangent line from the power available line at zero airspeed to the power

required curve. The tangent yields the best rate of climb for the least velocity or the most

vertical distance traveled for the least horizontal distance traveled, because it identifies the point

at which the most excess power occurs relative to the true airspeed. Typically, best rate of climb

is denoted by VY and best angle of climb by VX. Both are affected by altitude and gross weight

variations due to their association with the power required curve. An increase in weight shifts

the power required curve up and to the right so that the VX airspeed increases as the aircraft gets

heavier. It also happens that VX tends to be about 3/4 of VY.


7-28 PERFORMANCE

Figure 7-24 Best Angle of Climb

Figure 7-25 Rate of Climb vs. Best Angle of Climb

FLIGHT PHENOMENA 8-1

CHAPTER EIGHT

FLIGHT PHENOMENA

800. INTRODUCTION

This lesson introduces various phenomena associated with helicopter flight.













802. GENERAL

Maneuvering flight can place both the aircraft and the pilot under stress. Knowing the

maneuvering limitations is critical. In combat, the aircraft may be flown on the edge of the

envelope, as dictated by the mission or to save lives. In training, the student will be introduced

to the envelope gradually. Develop a comfort zone. Learning the dangers associated with flight

discussed in this chapter, as well as the methods for their avoidance/recovery, are critical to a

career in aviation.

803. FLIGHT ENVELOPE / V-N DIAGRAM

A helicopter’s design is dictated by the expected use. Performance and load bearing

requirements are set by the customer in the case of military aircraft or by the FAA and the

engineering and sales departments for civil aircraft. Anticipated strength requirements to meet

design criteria at a range of speeds are consolidated in a V-n diagram.

CHAPTER EIGHT HELICOPTER AERODYNAMICS WORKBOOK

8-2 FLIGHT PHENOMENA

The V-n diagram or V-G diagram is a graph that summarizes an aircraft's structural and

aerodynamic limitations at a particular weight, altitude and configuration. The horizontal axis is

indicated airspeed. The vertical axis of the graph is load factor, or G's. V-n diagrams define the

maneuvering envelope for fixed-wing aircraft and rotary-wing aircraft. Helicopter NATOPS

manuals typically do not have V-n diagrams because most helicopters do not have G-meters,

therefore pilots are unable to gauge loading. Rather, AOB limitations are developed with

consideration for associated load factors and general maneuver restrictions keep the aircraft in

the envelope.

Figure 8-1 V-n Diagram for Fixed Wing Aircraft

Several critical factors are identified on the V-n diagram. Even though NATOPS typically does

not include a V-n diagram, consideration of the following factors goes into development of those

maneuver limits and are worth knowing about before flying in critical situations:

Limit Load Factor. The top and bottom of the V-n diagram are established by the structural

limit line, or limit load factor. Limit load factor is the greatest load factor an aircraft can sustain

without any risk of permanent deformation. It is the maximum load factor anticipated in normal

daily operations. If the limit load factor is exceeded, some structural damage or permanent

deformation may occur. Aircraft will have both positive and negative limit load factors.

Overstress/Over-G is the condition of possible permanent deformation or damage that results

from exceeding the limit load factor. This type of damage will reduce the service life of the

aircraft because it weakens the aircraft's basic structure. Overstress/over-G may occur without

visibly damaging the airframe. Inside the aircraft are a variety of components, such as hydraulic

actuators and engine mounts, which are not designed to withstand the same loads that the

airframe can.

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER EIGHT


Figure 8-2 AH-64 Apache V-n Diagram

Ultimate load factor. Ultimate load factor is the maximum load factor that the aircraft can

withstand without structural failure. There will be some permanent deformation at the ultimate

load factor, but no actual failure of the major load-carrying components should occur. If you

exceed the ultimate load factor, structural failure is imminent (something major on the aircraft

will break). The ultimate load factor is 150% of the limit load factor.

Increases in gross weight and altitude require increases in AOA and lifting forces so that the

G envelope is reduced due to increased structural bending and blade flapping limits.

So far, in the roughly 70 years that helicopters have been operating, no really high load-factor

maneuver has been identified as a prohibitively hard to attain design consideration. Because

rotor blades are attached to the aircraft with a hinge or a relatively soft blade root, and because

centrifugal force tends to bend the blade down, a rotor blade flaps up but doesn’t bend

significantly when developing high thrust. As a consequence, the pilot need not worry about

causing a permanent set while doing an extreme maneuver. This is not to say, however, that no

structural damage has been done. Experience shows that high load-factor maneuvers raise the

level of oscillatory loads in the blades, hub, control system, and rotor-support structure. Usually

the pilot has a sense of these loads in the level of vibration they feel. Depending on the design of

the various components, the higher-than-normal oscillatory loads may cause fatigue damage that

shortens the useful life of the part.

Lift Limit. The left-hand side of the V-n diagram is the lift limit. This is the maximum load

factor available at a given airspeed. An aerodynamic limit of rotor thrust exists at speeds less

than maneuver speed because air mass flow through the rotor is decreasing and the rotor can't



generate high transient thrust levels. Attempted hard aft cyclic maneuvering at low speeds will

result in further reduced speed and "G available." Many helicopters have the capability of

generating in excess of four transient G's at high speed; it is unusual for a helicopter to be able to

sustain more than two G's.

Limit Airspeed. The vertical line on the right side is called the redline airspeed, or VNE

(Velocity never-to-exceed). Redline airspeed is the highest airspeed that an aircraft is allowed to

fly. Flight at speeds above VNE can cause structural damage. VNE is determined primarily by

excessive structural loads and power available, but may also be affected by controllability

limits, MCRIT, or airframe temperature.

Excessive structural loads may be encountered on components other than the main structural

members. Control surfaces, stabilizers, and other external components are often not able to

withstand the same forces that the wings (rotor disc) or fuselage can withstand. Maneuvering at

very high airspeeds may create sufficient forces to twist or break at their attachment point.

If an aircraft or component (advancing blade) reaches its critical Mach number (MCRIT) and is

not designed to withstand supersonic airflow, the shock waves generated may damage the

structure of the aircraft. Redline airspeed for these aircraft will be slightly below the airspeed at

which they will achieve MCRIT.

Redline airspeed may also be used to set limits on airframe temperature. As airspeed increases,

the aircraft encounters more air particles producing friction which heats up the airframe. This

heating can be extreme and hazardous at high speeds. Once the temperature becomes excessive,

the airframe may suffer creep damage.

Controllability may determine the redline airspeed on aircraft with conventional control systems.

At high airspeeds, dynamic pressure may create forces on the control surfaces which exceed the

pilot's ability to overcome. Or, due to the aeroelasticity of the control surfaces, full deflection of

the cockpit controls may cause only small deflection of the control surfaces. In either case, the

pilot will be unable to provide sufficient control input to safely maneuver the aircraft.

In fixed wing aircraft, the never exceed airspeed is established to preclude structural damage

from flutter. In a helicopter, never exceed speed is based upon power available or component

wear considerations.

VAFT. The maximum allowable rearward speed. This may be a structural limit, but

rotor/airframe configuration and rearward visibility from the cockpit are also factors, and VAFT

may be made as high as it is thought safe to test for.

Maneuvering Speed. The intersection of lift limit and structural limit lines occurs at the

maneuvering airspeed. Maneuvering airspeed (Va), also known as the corner airspeed, is the

maximum speed at which full control deflection can be abruptly applied without

overstressing the aircraft. Va varies with aircraft weight, just as the size of the maneuvering

envelope changes with weight. Above maneuver speed, the rotor can generate high aerodynamic

loads in excess of the limit load or can be pushed into retreating blade stall. Below maneuver



speed an aerodynamic limit of rotor thrust exists. At the maneuver speed the aircraft can achieve

limit load at a low speed, so it offers maximum turn rate and minimum turn radius.

Gust loading. Gust loading refers to the increase in the G load due to vertical wind gusts. The

load imposed by a gust is dependent upon the velocity of the gust. The higher the velocity, the

greater the increase in load. If an aircraft were generating the limit load factor during a

maximum performance turn and hit a vertical gust, the gust will instantaneously increase the

AOA of the airfoils and increase the lift on the rotor blades, enough to raise the G load above the

limit load factor. This is why "intentional flight through severe or extreme turbulence and

thunderstorms is prohibited" in many aircraft.

Vertical gusts of up to 30 feet per second may be encountered in moderate turbulence. This

could produce up to two G's of acceleration on the aircraft. Because gust loading is cumulative

with pilot-induced loading, the limit load factor due to pilot-induced loads should be reduced

to two-thirds of the normal limit load factor. For this reason, if you make the mistake of

entering a thunderstorm, consideration should be given to continuing to the other side since

maneuvering increases the pilot-induced loads.

Turbulence penetration also requires that you slow the aircraft to a speed that will reduce the

effects of stress caused by gust loading. Since a thunderstorm gives no assurance of positive G

loading, thunderstorm penetration airspeeds may reflect the intersection between the negative

load and negative lift lines.

804. VIBRATION ANALYSIS

Everything from your eyeballs to your aircraft has a natural frequency. This natural frequency is

determined by the components' mass and stiffness and is normally modeled as a spring mass

system in various modes of bending, and torsion.

Nodes are points where no motion occurs. As such, a node is a good place to suspend the rotor

system or locate crew or passenger seats for maximum comfort and minimum vibration.

The main source of vibrations for helicopters comes from the main rotor system. Vibrations

referred to as 1P, 2P, and 3P vibes are equal to the main rotor rotating frequency or multiples of

that frequency (vibrations per revolution). The frequency of the main rotor is a function of the

speed at which it rotates in revolutions per minute.

Figure 8-3, Vibration Analysis, provides a quick reference for basic analysis of vibrations

typically felt in the cockpit while flying. The number or beats of vibration related to the main

rotor blades can vary depending on the number of rotor blades installed, i.e., on an H-53E with

seven blades, several blades could be out of track rather than just one.



CATEGORY INDICATIONS CAUSE

LOW FREQUENCY (most

common)

1:1 LATERAL

1:1 VERTICAL

MRB out of balance

MRB out of track

2:1 Inherent in two-bladed helicopter.

Increase indicates worn rotating

control part of rotor hub part.

MEDIUM FREQUENCY 4:1 TO 6:1 Change in A/C ability to absorb

normal vibrations. Loose

component (landing gear most

common), loose cargo etc.

HIGH FREQUENCY TOO FAST TO COUNT

BUZZ IN PEDALS

Anything that rotates or vibrates at

the speed of the tail rotor

(transmission, engine, driveshaft).

Figure 8-3 Vibration Analysis

Tail shake. A problem that is usually worse in autorotation than in other flight conditions is

“tail shake.” This has been a significant problem on the prototypes of a number of helicopter

designs during their first test flights.

It is usually traced to unsteady airflow that arises at the main-rotor pylon or at the rotor hub, and

reaches the position of the tail rotor or empennage surfaces with high turbulence. If the

frequency of the turbulence happens to match one of the empennage’s natural frequencies, the

resulting resonance causes vibrations that can be felt throughout the entire helicopter.

The usual cure for this is to install special pylon fairings that act as low aspect ratio wings.

These produce tip vortices that tend to organize the flow and lower the turbulence downstream.

The unsteady flow from the hub can be suppressed by the installation of a round cap or “beanie”

that also produces vortices. Neither of these fixes should be done unless flight test results show

that they are necessary, since both add weight and drag to the helicopter.

Sometimes even these changes are not sufficient and it is necessary to avoid resonance by adding

weights which lower the natural frequencies of the vertical or horizontal stabilizer structure, or

raise the natural frequency with structural stiffening.

Other sources of vibrations due to external loads or airflows through the rotor system can also

cause excitement of component natural frequencies. The vibration caused by an oscillating

external load has at numerous times forced aircrew to pickle the load or, in a worse case, caused

the crash of an Israeli CH-53D that killed several personnel on board.

805. GROUND VORTEX

Occasionally during a discussion of takeoff operations, pilots will hear the term “ground vortex”

mentioned. As previously discussed, in a hover the rotor downwash travels outward from the

aircraft after impacting the ground. The height of this outward traveling airflow is

approximately equal to 1/3 the rotor diameter and has a curling tendency. This is called the



Ground Vortex. The speed of the vortex as it moves further from the aircraft slows due to

friction from the ground. As the helicopter moves forward, it catches up with the ground vortex,

and the rotor downwash mixes with increased relative wind to create a rotating vortex, which

eventually causes an increased downwash through the rotor system. This simulates a climbing

situation, increasing power required. Eventually this vortex is overrun at a higher speed.

Figure 8-4 Ground Vortex

806. COMPRESSIBILITY

Because the forward speed of the helicopter is added to the rotational velocity of the advancing

blade, the highest relative wind velocities occur at the tip of the advancing blade. When the

Mach number of the tip section of the advancing blade exceeds the Critical Mach number for

the rotor blade section, the results are compressibility effects. The principal effect of

compressibility is a large increase in drag and rearward shift of the airfoil aerodynamic center

(AC). Compressibility effects on the helicopter increase the power required to maintain rotor

RPM and cause rotor roughness, vibration, cyclic shake, and an undesirable structural twisting of

the rotor blade. Compressibility effects become more severe at higher lift coefficients (higher

blade angles of attack) and higher Mach numbers.

Adverse Compressibility Conditions. The following operating conditions represent the

conditions that contribute to compressibility:

1. High airspeed.



2. High rotor RPM.

3. High gross weight.

4. High DA.

5. High-G maneuvers.

6. Low temperature—the speed of sound is proportional to the square root of the absolute

temperature; therefore, it decreases as temperature decreases.

7. Turbulent air—sharp gusts momentarily increase the blade AOA and thus lower the

Critical Mach number to the point where compressibility effects may be encountered on the

blade.

Corrective Actions. Corrective actions are any actions that will decrease the AOA or velocity

of the airflow. There are similarities in the critical conditions for compressibility and retreating

blade stall, with one notable exception: compressibility occurs at high rotor RPM, and retreating

blade stall occurs at low rotor RPM. With the exception of RPM control, the recovery technique

is identical for both. Such techniques include decreasing:

1. Blade pitch by lowering collective, if possible.

2. Rotor RPM.

3. Severity of maneuver.

4. Airspeed.

Critical Mach Number. Critical Mach Number is the flow speed at which the local velocity at

some point on an airfoil first reaches sonic speed. Because airfoils speed up flow on the upper

surface to generate lift, the flow over the top is faster than the free stream. When the free stream

past a section of the rotor blade is going about Mach .72, at the point of maximum velocity over

the airfoil’s surface the local speed may reach Mach 1.0, or the speed of sound (Figure 10-25).

That flow Mach number, .72, is known as the Critical Mach Number. Actual Critical Mach

Number depends upon the shape of the airfoil.



Figure 8-5 Critical Mach Number

Drag Divergence Mach Number (MDD). Drag Divergence Mach is a speed that exists above

the critical Mach number but below sonic velocity. At speed above the Critical Mach Number,

air becomes more and more compressible. It begins to form a shock wave that increases drag

and disrupts flow. As the flow over the airfoil moves faster and faster, stronger shock waves

begin to form on the airfoil. The flow disruption and strong pressure disturbances greatly

increase drag and cause airflow separation. Drag due to compression starts out small at lower

speeds, but at some point before sonic velocity begins to dramatically increase. The Mach

number at which the drag dramatically increases is called Drag or Force Divergence Mach

Number.

Figure 8-6 Drag Increase with Mach Number

The highest speed encountered by the blades in forward flight occurs at a rotor blade’s tip as it

passes the 3 o’clock position. At that point the velocity equals the RPM times the blade radius,

plus the helicopter’s forward velocity. When velocity at the tip on the advancing side

approaches MDD an increase in power is required to overcome extra drag. The fact that drag

increases at one point on the rotor disk and not at others is felt as a vibration. Also, as

compressibility increases near MDD on the advancing blade, there is an increase in vibrations and



structural stress as the AC of the rotor blade migrates rearward in the transonic region. An

additional undesirable effect is noise as blades repeatedly “break the sound barrier” as they go

around the disc’s advancing side.

The solutions typically used to deal with advancing blade compressibility effects are sweeping

the leading edge of the rotor blades back, varying the airfoil thickness along the span, and

varying the airfoil section along the span. Sweep reduces the velocity that the blade tip "sees"

thereby delaying drag divergence and reduces the CL max of the airfoil.

Variation of airfoil thickness and of airfoil section serve to change the properties of the airfoil.

As the rotational velocity increases out towards the end of the blade, the thickness decreases or

the overall qualities of the airfoil change to take advantage of the increase in speed.

807. RETREATING BLADE STALL (RBS)

In forward flight, the relative airflow through the main rotor disk is different on the advancing

and retreating side. The relative airflow over the advancing side is higher due to the forward

speed of the helicopter, while the relative airflow on the retreating side is lower. This

dissymmetry of lift increases as forward speed increases.

To generate the same amount of lift across the rotor disk, the advancing blade flaps up while the

retreating blade flaps down. This causes the AOA to decrease on the advancing blade, which

reduces lift, and increase on the retreating blade, which increases lift. At some point as the

forward speed increases, the low blade speed on the retreating blade, and its high AOA cause a

stall and loss of lift.

Retreating blade stall is a factor in limiting a helicopter’s never-exceed speed (VNE ) and its

development can be felt by a low frequency vibration, pitching up of the nose, and a roll in the

direction of the retreating blade. High weight, low rotor RPM, high density altitude, turbulence

and/or steep, abrupt turns are all conducive to retreating blade stall at high forward airspeeds. As

altitude is increased, higher blade angles are required to maintain lift at a given airspeed. Thus,

retreating blade stall is encountered at a lower forward airspeed at altitude. Most manufacturers

publish charts and graphs showing a VNE decrease with altitude.

When recovering from a retreating blade stall condition caused by high airspeed, moving the

cyclic aft only worsens the stall as aft cyclic produces a flare effect, thus increasing the AOA.

Pushing forward on the cyclic also deepens the stall as the AOA on the retreating blade is

increased. While the first step in a proper recovery is usually to reduce collective, RBS should

be evaluated in light of the relevant factors discussed in the previous paragraph and addressed

accordingly. For example, if a pilot at high weight and high DA is about to conduct a high

reconnaissance prior to a confined area operation where rolling into a steep turn causes onset of

RBS, the recovery is to roll out of the turn. If the cause is low rotor RPM, then increase the

RPM.



808. GROUND RESONANCE

Helicopters with articulating rotors (usually designs with three or more main rotor blades) are

subject to ground resonance, a destructive vibration phenomenon that occurs at certain rotor

speeds when the helicopter is on the ground. Ground resonance is a mechanical design issue that

results from the helicopter’s airframe having a natural frequency that can be intensified by an

out-of-balance rotor. The unbalanced rotor disk vibrates at the same frequency (or multiple

thereof) of the airframe’s resonant frequency, and the harmonic oscillation increases because the

engine is adding power to the system, increasing the magnitude (amplitude) of the vibrations

until the structure or structures fail. This condition can cause a helicopter to self-destruct in a

matter of seconds.

Hard contact with the ground on one corner (and usually with wheel-type landing gear) can send

a shockwave to the main rotor head, resulting in the blades of a three-blade rotor disk moving

from their normal 120° relationship to each other. This movement occurs along the drag hinge

and could result in something like 122°, 122°, and 116° between blades. When another part of

the landing gear strikes the surface, the unbalanced condition could be further aggravated.

If the RPM is low, the only corrective action to stop ground resonance is to close the throttle

immediately and fully lower the collective to place the blades in low pitch. If the RPM is in the

normal operating range, fly the helicopter off the ground, and allow the blades to re-phase

themselves automatically. Then, make a normal touchdown. If a pilot lifts off and allows the

helicopter to firmly re-contact the surface before the blades are realigned, a second shock could

move the blades again and aggravate the already unbalanced condition. This could lead to a

violent, uncontrollable oscillation.

This situation does not occur in rigid or semi-rigid rotor disks because there is no drag hinge. In

addition, skid-type landing gear is not as prone to ground resonance as wheel-type landing gear,

since the rubber tires' resonant frequency typically can match that of the spinning rotor, unlike

the condition of a rigid landing gear.

809. DYNAMIC ROLLOVER

A helicopter is susceptible to a lateral rolling tendency, called dynamic rollover, when it is in

contact with the surface during takeoffs or landings. For dynamic rollover to occur, some factor

must first cause the helicopter to roll or pivot around a skid or landing gear wheel, until its

critical rollover angle is reached. The angle at which dynamic rollover occurs will vary based on

helicopter type. Then, beyond this point, main rotor thrust continues the roll and recovery is

impossible. After this angle is achieved, the cyclic does not have sufficient range of control to

eliminate the thrust component and convert it to lift. If the critical rollover angle is exceeded, the

helicopter rolls on its side regardless of the cyclic corrections made.

Dynamic rollover begins when the helicopter starts to pivot laterally around its skid or wheel.

For dynamic rollover to occur the following three factors must be present:

1. A rolling moment



2. A pivot point other than the helicopter’s normal CG

3. Thrust greater than weight

This can occur for a variety of reasons, including the failure to remove a tie down or skid-

securing device, or if the skid or wheel contacts a fixed object while hovering sideward, or if the

gear is stuck in ice, soft asphalt, or mud. Dynamic rollover may also occur if you use an

improper landing or takeoff technique or while performing slope operations. Whatever the

cause, dynamic rollover is possible if not using the proper corrective technique.

Once started, dynamic rollover cannot be stopped by application of opposite cyclic control alone.

For example, the right skid contacts an object and becomes the pivot point while the helicopter

starts rolling to the right. Even with full left cyclic applied, the main rotor thrust vector and its

moment follows the aircraft as it continues rolling to the right. Quickly reducing collective pitch

is the most effective way to stop dynamic rollover from developing. Dynamic rollover can occur

with any type of landing gear and all types of rotor disks.

It is important to remember rotor blades have a limited range of movement. If the tilt or roll of

the helicopter exceeds that range (5–8°), the controls (cyclic) can no longer command a vertical

lift component and the thrust or lift becomes a lateral force that rolls the helicopter over. When

limited rotor blade movement is coupled with the fact that most of a helicopter’s weight is high

in the airframe, another element of risk is added to an already slightly unstable center of gravity.

Pilots must remember that in order to remove thrust, the collective must be lowered as this is the

only recovery technique available.

Critical Conditions

Certain conditions reduce the critical rollover angle, thus increasing the possibility for dynamic

rollover and reducing the chance for recovery. The rate of rolling motion is also a consideration

because, as the roll rate increases, there is a reduction of the critical rollover angle at which

recovery is still possible. Other critical conditions include operating at high gross weights with

thrust (lift) approximately equal to the weight.

1. The following conditions are most critical for helicopters with counterclockwise rotor

rotation:

2. Right side skid or landing wheel down, since translating tendency adds to the rollover

force.

3. Right lateral center of gravity (CG).

4. Crosswinds from the left.

5. Left yaw inputs.

For helicopters with clockwise rotor rotation, the opposite conditions would be true.



Figure 8-7 Dynamic rollover

Cyclic Trim

When maneuvering with one skid or wheel on the ground, care must be taken to keep the

helicopter cyclic control carefully adjusted. For example, if a slow takeoff is attempted and the

cyclic is not positioned and adjusted to account for translating tendency, the critical recovery

angle may be exceeded in less than two seconds. Control can be maintained if the pilot

maintains proper cyclic position and does not allow the helicopter’s roll and pitch rates to

become too great. Fly the helicopter into the air smoothly while keeping movements of pitch,

roll, and yaw small; do not allow any abrupt cyclic pressures.

Normal Takeoffs and Landings

Dynamic rollover is possible even during normal takeoffs and landings on relatively level

ground, if one wheel or skid is on the ground and thrust (lift) is approximately equal to the

weight of the helicopter. If the takeoff or landing is not performed properly, a roll rate could

develop around the wheel or skid that is on the ground. When taking off or landing, perform the

maneuver smoothly and carefully adjust the cyclic so that no pitch or roll movement rates build

up, especially the roll rate. If the bank angle starts to increase to an angle of approximately 5–8°,

and full corrective cyclic does not reduce the angle, the collective should be reduced to diminish

the unstable rolling condition. Excessive bank angles can also be caused by landing gear caught

in a tie down strap, or a tie down strap still attached to one side of the helicopter. Lateral loading

imbalance (usually outside published limits) is another contributing factor.



Slope Takeoffs and Landings

During slope operations, excessive application of cyclic control into the slope, together with

excessive collective pitch control, can result in the downslope skid or landing wheel rising

sufficiently to exceed lateral cyclic control limits, and an upslope rolling motion can occur.

When performing slope takeoff and landing maneuvers, follow the published procedures and

keep the roll rates small. Slowly raise the downslope skid or wheel to bring the helicopter level,

and then lift off. During landing, first touchdown on the upslope skid or wheel, then slowly

lower the downslope skid or wheel using combined movements of cyclic and collective. If the

helicopter rolls approximately 5–8° to the upslope side, decrease collective to correct the bank

angle and return to level attitude, then start the landing procedure again.

Figure 8-8 Slope takeoff or landing 1

Use of Collective

The collective is more effective in controlling the rolling motion than lateral cyclic, because it

reduces the main rotor thrust (lift). A smooth, moderate collective reduction, at a rate of less

than approximately full up to full down in two seconds, may be adequate to stop the rolling

motion. Take care, therefore, not to dump collective at an excessively high rate, as this may

cause a main rotor blade to strike the fuselage. Additionally, if the helicopter is on a slope and

the roll starts toward the upslope side, reducing collective too fast may create a high roll rate in

the opposite direction. When the upslope skid or wheel hits the ground, the dynamics of the

motion can cause the helicopter to bounce off the upslope skid or wheel, and the inertia can

cause the helicopter to roll about the downslope ground contact point and over on its side.

Under normal conditions on a slope, the collective should not be pulled suddenly to get airborne

because a large and abrupt rolling moment in the opposite direction could occur. Excessive



application of collective can result in the upslope skid or wheel rising sufficiently to exceed

lateral cyclic control limits. This movement may be uncontrollable. If the helicopter develops a

roll rate with one skid or wheel on the ground, the helicopter can roll over on its side.

Figure 8-9 Slope takeoff or landing 2

Precautions

To help avoid dynamic rollover:

1. Always practice hovering autorotations into the wind, and be wary when the wind is gusty

or greater than 10 knots.

2. Use extreme caution when hovering close to fences, sprinklers, bushes, runway/taxi lights,

tie-down cables, deck nets, or other obstacles that could catch a skid or wheel. Aircraft parked

on hot asphalt overnight might find the landing gear sunk in and stuck as the ramp cooled during

the evening.

3. Always use a two-step lift-off. Pull in just enough collective pitch control to be light on the

skids or landing wheels and feel for equilibrium, then gently lift the helicopter into the air.

4. Hover high enough to have adequate skid or landing wheel clearance from any obstacles

when practicing hovering maneuvers close to the ground, especially when practicing sideways or

rearward flight.

5. Remember that when the wind is coming from the upslope direction, less lateral cyclic

control is available.

6. Avoid tailwind conditions when conducting slope operations.



7. Remember that less lateral cyclic control is available due to the translating tendency of the

tail rotor when the left skid or landing wheel is upslope. (This is true for counterclockwise rotor

disks.)

8. Keep in mind that the lateral cyclic requirement changes when passengers or cargo are

loaded or unloaded.

9. Be aware that if the helicopter utilizes interconnecting fuel lines that allow fuel to

automatically transfer from one side of the helicopter to the other, the gravitational flow of fuel

to the downslope tank could change the CG, resulting in a different amount of cyclic control

application to obtain the same lateral result.

10. Do not allow the cyclic limits to be reached. If the cyclic control limit is reached, further

lowering of the collective may cause mast bumping. If this occurs, return to a hover and select a

landing point with a lesser degree of slope.

11. During a takeoff from a slope, begin by leveling the main rotor disk with the horizon or

very slightly into the slope to ensure vertical lift and only enough lateral thrust to prevent sliding

on the slope. If the upslope skid or wheel starts to leave the ground before the downslope skid or

wheel, smoothly and gently lower the collective and check to see if the downslope skid or wheel

is caught on something. Under these conditions, vertical ascent is the only acceptable method of

lift-off.

12. Be aware that dynamic rollover can be experienced during flight operations on a floating

platform if the platform is pitching/rolling while attempting to land or takeoff. Generally, the

pilot operating on floating platforms (barges, ships, etc.) observes a cycle of seven during which

the waves increase and then decrease to a minimum. It is that time of minimum wave motion

that the pilot needs to use for the moment of landing or takeoff on floating platforms. Pilots

operating from floating platforms should also exercise great caution concerning cranes, masts,

nearby boats (tugs) and nets.

810. LOW-G CONDITIONS

“G” is an abbreviation for acceleration due to the earth’s gravity. A person standing on the

ground or sitting in an aircraft in level flight is experiencing one G. An aircraft in a tight, banked

turn with the pilot being pressed into the seat is experiencing more than one G or high-G

conditions. A person beginning a downward ride in an elevator or riding down a steep track on a

roller coaster is experiencing less than one G or low-G conditions. The best way for a pilot to

recognize low G is a weightless feeling similar to the start of a downward elevator ride.

Helicopters rely on positive G to provide much or all of their response to pilot control inputs.

The pilot uses the cyclic to tilt the rotor disk, and, at one G, the rotor is producing thrust equal to

aircraft weight. The tilting of the thrust vector provides a moment about the center of gravity to

pitch or roll the fuselage. In a low-G condition, the thrust and consequently the control authority

are greatly reduced.



Although their control ability is reduced, multi-bladed (three or more blades) helicopters can

generate some moment about the fuselage independent of thrust due to the rotor hub design with

the blade attachment offset from the center of rotation. However, helicopters with two-bladed

teetering rotors rely entirely on the tilt of the thrust vector for control. Therefore, low-G

conditions can be catastrophic for two-bladed helicopters.

At lower speeds, such as initiation of a takeoff from hover or the traditional recovery from vortex

ring state, forward cyclic maneuvers do not cause low G and are safe to perform. However, an

abrupt forward cyclic input or pushover in a two-bladed helicopter can be dangerous and must be

avoided, particularly at higher speeds. During a pushover from moderate or high airspeed, as the

helicopter noses over, it enters a low-G condition. Thrust is reduced, and the pilot has lost

control of fuselage attitude but may not immediately realize it. Tail rotor thrust or other

aerodynamic factors will often induce a roll. The pilot still has control of the rotor disk, and may

instinctively try to correct the roll, but the fuselage does not respond due to the lack of thrust. If

the fuselage is rolling right, and the pilot puts in left cyclic to correct, the combination of

fuselage angle to the right and rotor disk angle to the left becomes quite large and may exceed

the clearances built into the rotor hub. This results in the hub contacting the rotor mast, which is

known as mast bumping. Low-G mast bumping has been the cause of numerous military and

civilian fatal accidents. It was initially encountered during nap-of-the-earth flying, a very low-

altitude tactical flight technique used by the military where the aircraft flies following the

contours of the geographical terrain. The accident sequence may be extremely rapid, and the

energy and inertia in the rotor system can sever the mast or allow rotor blades to strike the tail or

other portions of the helicopter.

Turbulence, especially severe downdrafts, can also cause a low-G condition and, when combined

with high airspeed, may lead to mast bumping. Typically, helicopters handle turbulence better

than a light airplane due to smaller surface area of the rotor blades. During flight in turbulence,

momentary excursions in airspeed, altitude, and attitude are to be expected. Pilots should

respond with smooth, gentle control inputs and avoid over controlling. Most importantly, pilots

should slow down, as mast bumping is less likely at lower airspeeds.

Multi-bladed rotors may experience a phenomenon similar to mast bumping known as droop

stop pounding if flapping clearances are exceeded, but because they retain some control authority

at low G, occurrences are less common than for teetering rotors.

811. LOW ROTOR RPM AND ROTOR STALL

Rotor RPM is a critically important parameter for all helicopter operations. Just as airplanes will

not fly below a certain airspeed, helicopters will not fly below a certain rotor RPM. Safe rotor

RPM ranges are marked on the helicopter’s tachometer and specified in the Rotorcraft Flight

Manual (RFM). If the pilot allows the rotor RPM to fall below the safe operating range, the

helicopter is in a low RPM situation. If the rotor RPM continues to fall, the rotor will eventually

stall.

Rotor stall should not be confused with retreating blade stall, which occurs at high forward

speeds and over a small portion of the retreating blade tip. Retreating blade stall causes vibration



and control problems, but the rotor is still very capable of providing sufficient lift to support the

weight of the helicopter. Rotor stall, however, can occur at any airspeed, and the rotor quickly

stops producing enough lift to support the helicopter, causing it to lose lift and descend rapidly.

Rotor stall is very similar to the stall of an airplane wing at low airspeeds. The airplane wing

relies on airspeed to produce the required airflow over the wing, whereas the helicopter relies on

rotor RPM. As the airspeed of the airplane decreases or the speed of the helicopter rotor slows

down, the AOA of the wing/rotor blade must be increased to support the weight of the aircraft.

At a critical angle (about 15°), the airflow over the wing or the rotor blade will separate and stall,

causing a sudden loss of lift and increase in drag. An airplane pilot recovers from a stall by

lowering the nose to reduce the AOA and adding power to restore normal airflow over the wing.

However, the falling helicopter is experiencing upward airflow through the rotor disk, and the

resulting AOA is so high that even full down collective will not restore normal airflow. In the

helicopter when the rotor stalls, it does not do so symmetrically because any forward airspeed

will produce a higher airflow on the advancing side than on the retreating side. This causes the

retreating blade to stall first, and its weight makes it descend as it moves aft while the advancing

blade is climbing as it goes forward. The resulting low aft blade and high forward blade become

a rapid aft tilting of the rotor disc sometimes referred to as rotor “blow back” or “flap back.” As

the helicopter begins to descend, the upward flow of air acting on the bottom surfaces of the tail

boom and any horizontal stabilizers tend to pitch the aircraft nose down. These two effects,

combined with any aft cyclic by the pilot attempting to keep the aircraft level, allow the rotor

blades to blow back and contact the tail boom, in some cases actually severing the tail boom.

Since the tail rotor is geared to the main rotor, in many helicopters the loss of main rotor RPM

also causes a significant loss of tail rotor thrust and a corresponding loss of directional control.

Rotor stalls in helicopters are not recoverable. At low altitude, rotor stall will result in an

accident with significant damage to the helicopter, and at altitudes above approximately 50 feet

the accident will likely be fatal. Consequently, early recognition of the low rotor RPM condition

and proper recovery technique is imperative.

Low rotor RPM can occur during power-off and power-on operations. During power-off flight, a

low RPM situation can be caused by the failure to quickly lower the collective after an engine

failure or by raising the collective at too great a height above ground at the bottom of an

autorotation. However, more common are power-on rotor stall accidents. These occur when the

engine is operating normally but the pilot demands more power than is available by pulling up

too much on the collective. Known as “overpitching,” this can easily occur at higher density

altitudes where the engine is already producing its maximum horsepower and the pilot raises the

collective. The corresponding increased AOA of the blades requires more engine horsepower to

maintain the speed of the blades; however, the engine cannot produce any additional horsepower,

so the speed of the blades decreases. A similar situation can occur with a heavily loaded

helicopter taking off from a confined area. Other causes of a power-on low rotor RPM condition

include the pilot rolling the throttle the wrong way in helicopters not equipped with a governor or

a governor failure in helicopters so equipped.

As the RPM decreases, the amount of horsepower the engine can produce also decreases. Engine

horsepower is directly proportional to its RPM, so a 10 percent loss in RPM due to overpitching,



or one of the other scenarios above, will result in a 10 percent loss in the engine’s ability to

produce horsepower, making recovery even slower and more difficult than it would otherwise

be. With less power from the engine and less lift from the decaying rotor RPM, the helicopter

will start to settle. If the pilot raises the collective to stop the settling, the situation will feed

upon itself rapidly leading to rotor stall.

There are a number of ways the pilot can recognize the low rotor RPM situation. Visually, the

pilot can not only see the rotor RPM indicator decrease but also the change in torque will

produce a yaw; there will also be a noticeable decrease in engine noise, and at higher airspeeds

or in turns, an increase in vibration. Many helicopters have a low RPM warning system that

alerts the pilot to the low rotor RPM condition.

To recover from the low rotor RPM condition the pilot must simultaneously lower the collective,

increase throttle if available and apply aft cyclic to maintain a level attitude. At higher airspeeds,

additional aft cyclic may be used to help recover lost RPM. Recovery should be accomplished

immediately before investigating the problem and must be practiced to become a conditioned

reflex.



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TAIL ROTOR CONSIDERATIONS 9-1

CHAPTER NINE

TAIL ROTOR CONSIDERATIONS

900. INTRODUCTION

The purpose of this chapter is to aid the student in understanding tail rotor design. Tail rotor

design considerations include both performance and rotor configuration requirements.












902. TORQUE EFFECT

For purposes of uniformity, conventional main rotor direction is chosen to be counterclockwise

as viewed from above. A single main rotor imparts a moment on the fuselage which, if left

unbalanced, would cause the fuselage to rotate clockwise around the vertical axis. This moment

is compensated for by placing an anti-torque tail rotor a certain distance from the center of

gravity of the aircraft. The thrust of the anti-torque tail rotor multiplied by the distance to the

CG results in a moment in the opposite direction to that generated by the main rotor. If the

forces and moments involved are considered in combination, the moments balance, but the

forces do not. The unbalanced force of the tail rotor causes a right translating tendency that is

most noticeable in a hover and occurs to a lesser extent in forward flight. Additionally, any

change in power setting will change the torque and therefore yaw. The effects of wind on the tail

rotor’s effectiveness must also be considered if the helicopter is to be usable in a wide range of

operating conditions.

CHAPTER NINE HELICOPTER AERODYNAMICS WORKBOOK

9-2 TAIL ROTOR CONSIDERATIONS

Figure 9-1 Tail Rotor Unbalanced Force

Pilots of helicopters with a clockwise rotation (and thus a left main rotor moment) simply need to

consider the effects in the opposite direction.

Figure 6-2 below illustrates the method of torque balance and directional control for both the tail

rotor configuration and alternate methods.

Helicopter Configuration

Torque Balance Directional Control

Yaw Moment

Single MR, TR TR Thrust TR Collective

Coaxial MR diff torque MR diff collective

Tandem MR diff torque MR diff cyclic

Side-by-side MR diff torque MR diff cyclic

Figure 9-2 Yaw Control Mechanisms for Various Configurations

903. VERTICAL STABILIZER

A vertical stabilizer can help quite a bit in reducing the amount of tail rotor thrust required in

forward flight. Shaped like a wing, a vertical stabilizer provides lift (thrust) in the direction of

anti-torque. The vertical stabilizer can be either a cambered airfoil or a symmetrical airfoil

mounted on an offset angle. The higher the aircraft’s velocity, the more the vertical stabilizer

will be contributing to the anti-torque effort. At higher speeds, tail rotor power requirements are

significantly reduced, therefore more engine power is now available to drive the main rotor

system.

HELICOPTER AERODYNAMICS WORKBOOK CHAPTER NINE


Design tradeoffs have precluded the production of any military helicopters that can actually fly

in level, balanced flight with a complete tail rotor failure because of the power interactions

between the tail rotor and vertical fin. Making a large enough tail fin that could completely

compensate for a lost tail rotor would compromise sideward flight capability. The Apache and

Blackhawk, however, were designed to be able to fly straight in a controlled descent (at a

specified airspeed) without the tail rotor operating.

Figure 9-3 Vertical Stabilizer

904. TRANSLATING TENDENCY AND HOVER ATTITUDE

While the tail rotor system produces anti-torque effect, it also produces thrust in the horizontal

plane, causing the aircraft to drift right laterally in a hover, for a counterclockwise rotating,

single-rotor helicopter. The aviator must compensate for this right translating tendency of the

helicopter by tilting the main rotor disk to the left. This lateral tilt creates an equal but opposite

main rotor force to the left that compensates for the tail rotor thrust to the right. These two

horizontal forces, however, are often offset from each other vertically. The main rotor force to

the left coupled with the tail rotor force to the right commonly causes a left skid low hover

attitude during flight.



Figure 9-4 Translating Tendency

905. WEATHER VANING

In a no-wind hover, the tail rotor provides all of the anti-torque compensation. As the aircraft

moves into forward flight, the tail rotor is assisted in this compensatory effort by the weather-

vaning effect and the vertical stabilizer. The increased parasitic drag produced on the

longitudinal surface of the aircraft as the relative wind increases causes the aircraft to "steer" into

the relative wind. This weather-vaning effect will increase proportionally with airspeed and

provide minor assistance to the anti-torque effect.

Figure 9-5 Weather Vaning

906. TAIL ROTOR FAILURES AND ISSUES

Anti-torque malfunctions may occur through a number of mechanisms: a loss of the entire

gearbox/components; a fixed pedal setting, left, right, or neutral; driveshaft failure; loss of tail

rotor authority/loss of tail rotor effectiveness (LTA/LTE). Also, even though an engine failure

removes the need for anti-torque compensation, directional control at touchdown may be limited

in a number of situations.

1. LTA and LTE

The ability of the tail rotor to provide anti-torque and yaw control can be greatly reduced by two

factors that are easily confused. LTA is related to power available to the main and tail rotor.



LTE is related to the direction from which the wind strikes the tail rotor in a hover and tends to

be labeled as an aerodynamic phenomenon compared to LTA, which is most often described as a

mechanical phenomenon.

2. LTA

This occurs when power required for hover exceeds power available. Power supplied to the

main rotor is delivered as a torque at a certain RPM.

Power = Torque x RPM

When the engines are providing the maximum that they are capable of at 100% RPM it will

translate to a certain amount of torque. If the pilot demands more performance by continuing to

increase collective the AOA on the main rotor blades will increase. Lift will increase, but so will

drag. Because power is a constant, the main rotor response to the increased drag will be an

increase in torque and a decrease in RPM.

Tail rotor thrust required for flight is a function of main rotor torque. Tail rotor thrust available

is a function of RPM squared. When the main rotor slows down it also slows down the tail rotor,

providing less tail rotor thrust. Thus, the increased tail rotor thrust required to counteract

increasing main rotor torque with drooped turns is not available. The pilot can call for more

tail rotor thrust by increasing tail rotor torque with increased left pedal, but at some point the

ability to increase tail rotor AOA runs out. When tail rotor thrust required exceeds tail rotor

thrust available, LTA occurs and the nose of the aircraft yaws to the right.

3. LTE

Loss of tail rotor effectiveness (LTE) or an unanticipated yaw is defined as an un-commanded,

rapid yaw towards the advancing blade which does not subside of its own accord. It can result in

the loss of the aircraft if left unchecked. It is very important for pilots to understand that LTE is

caused by an aerodynamic interaction between the main rotor and tail rotor and not caused from

a mechanical failure. Some helicopter types are more likely to encounter LTE due to the normal

certification thrust produced by having a tail rotor that, although meeting certification standards,

is not always able to produce the additional thrust demanded by the pilot.

A helicopter is a collection of compromises. Compare the size of an airplane propeller to that of

a tail rotor. Then, consider the horsepower required to run the propeller. For example, a Cessna

172P is equipped with a 160-horsepower (HP) engine. A Robinson R-44 with a comparably

sized tail rotor is rated for a maximum of 245 HP. If you assume the tail rotor consumes 50 HP,

only 195 HP remains to drive the main rotor. If the pilot were to apply enough collective to

require 215 HP from the engine, and enough left pedal to require 50 HP for the tail rotor, the

resulting engine overload would lead to one of two outcomes: slow down (reduction in RPM) or

premature failure. In either outcome, anti-torque would be insufficient and total lift might be

less than needed to remain airborne.



Every helicopter design requires some type of anti-torque system to counteract main rotor torque

and prevent spinning once the helicopter lifts off the ground. A helicopter is heavy, and the

power-plant places a high demand on fuel. Weight penalizes performance, but all helicopters

must have an anti-torque system, which adds weight. Therefore, the tail rotor is certified for

normal flight conditions. Environmental forces can overwhelm any aircraft, rendering the

inherently unstable helicopter especially vulnerable.

As with any aerodynamic condition, it is very important for pilots to not only to understand the

definition of LTE, but more importantly, how and why it happens, how to avoid it, and lastly,

how to correct it once it is encountered. We must first understand the capabilities of the aircraft

or even better what it is not capable of doing. For example, if you were flying a helicopter with a

maximum gross weight of 5,200 lb., would you knowingly try to take on fuel, baggage and

passengers causing the weight to be 5,500 lb.? A wise professional pilot should not ever exceed

the certificated maximum gross weight or performance flight weight for any aircraft. The

manuals are written for safety and reliability. The limitations and emergency procedures are

stressed because lapses in procedures or exceeding limitations can result in aircraft damage or

human fatalities. At the very least, exceeding limitations will increase the costs of maintenance

and ownership of any aircraft and especially helicopters.

Overloaded parts may fail before their designed lifetime. There are no extra parts in helicopters.

The respect and discipline pilots exercise in following flight manuals should also be applied to

understanding aerodynamic conditions. If flight envelopes are exceeded, the end results can be

catastrophic.

LTE is an aerodynamic condition and is the result of a control margin deficiency in the tail rotor.

It can affect all single-rotor helicopters that utilize a tail rotor. The design of main and tail rotor

blades and the tail boom assembly can affect the characteristics and susceptibility of LTE but

will not nullify the phenomenon entirely. Translational lift is obtained by any amount of clean

air through the main rotor disk. The same holds true for the tail rotor. As the tail rotor works in

less turbulent air, it reaches a point of translational thrust. At this point, the tail rotor becomes

aerodynamically efficient and the improved efficiency produces more anti-torque thrust. The

pilot can determine when the tail rotor has reached translational thrust. As more anti-torque

thrust is produced, the nose of the helicopter yaws to the left (opposite direction of the tail rotor

thrust), forcing the pilot to correct with right pedal application (actually decreasing the left

pedal). This, in turn, decreases the AOA in the tail rotor blades. Pilots should be aware of the

characteristics of the helicopter they fly and be particularly aware of the amount of tail rotor

pedal typically required for different flight conditions.

LTE is a condition that occurs when the flow of air through a tail rotor is altered in some way, by

altering the angle or speed at which the air passes through the rotating blades of the tail rotor

disk. As discussed in the previous paragraph, an effective tail rotor relies on a stable and

relatively undisturbed airflow in order to provide a steady and constant anti-torque reaction. The

pitch and AOA of the individual blades will determine the thrust. A change to either of these

alters the amount of thrust generated. A pilot’s yaw pedal input causes a thrust reaction from the

tail rotor. Altering the amount of thrust delivered for the same yaw input creates an imbalance.

Taking this imbalance to the extreme will result in the loss of effective control in the yawing

plane, and LTE will occur.



This alteration of tail rotor thrust can be affected by numerous external factors. The main factors

contributing to LTE are:

a. Airflow and downdraft generated by the main rotor blades interfering with the airflow

entering the tail rotor assembly.

b. Main blade vortices developed at the main blade tips entering the tail rotor disk.

c. Turbulence and other natural phenomena affecting the airflow surrounding the tail

rotor.

d. A high-power setting, hence large main rotor pitch angle, induces considerable main

rotor blade downwash and hence more turbulence than when the helicopter is in a low

power condition.

e. A slow forward airspeed, typically at speeds where translational lift and translational

thrust are in the process of change and airflow around the tail rotor will vary in

direction and speed.

f. The airflow relative to the helicopter;

i. Worst case—relative wind within ±15° of the 10 o’clock position, generating

vortices that can blow directly into the tail rotor. This is dictated by the

characteristics of the helicopters aerodynamics of tail-boom position, tail rotor

size and position relative to the main rotor and vertical stabilizer, size and

shape.

ii. Weathercock stability—tailwinds from 120° to 240°, such as left crosswinds,

causing high pilot workload.

iii. Tail rotor vortex ring state (210° to 330°). Winds within this region will result

in the development of the vortex ring state of the tail rotor.

g. Combinations (i, ii, iii) of these factors in a particular situation can easily require

more anti-torque than the helicopter can generate and in a particular environment

LTE can be the result.

Certain flight activities lend themselves to being at higher risk of LTE than others. For example,

power line and pipeline patrol sectors, low speed aerial filming/photography as well as in the

Police and Helicopter Emergency Medical Services (EMS) environments can find themselves in

low-and-slow situations over geographical areas where the exact wind speed and direction are

hard to determine.

Unfortunately, the aerodynamic conditions that a helicopter is susceptible to are not explainable

in black and white terms. LTE is no exception. There are a number of contributing factors, but

what is more important in preventing LTE is to note them, and then to associate them with



situations that should be avoided. Whenever possible, pilots should learn to avoid the following

combinations:

a. Low and slow flight outside of ground effect.

b. Winds from ±15º of the 10 o’clock position and probably on around to 5 o’clock

position.

c. Tailwinds that may alter the onset of translational lift and translational thrust, and

hence induce high power demands and demand more anti-torque (left pedal) than the

tail rotor can produce.

d. Low speed downwind turns.

e. Large changes of power at low airspeeds.

f. Low speed flight in the proximity of physical obstructions that may alter a smooth

airflow to both the main rotor and tail rotor.

Pilots who put themselves in situations where the combinations above occur should know that

they are likely to encounter LTE. The key is not to put the helicopter in a compromising

condition, while at the same time being educated enough to recognize the onset of LTE and

being prepared to react quickly to it before the helicopter cannot be controlled.

Early detection of LTE, followed by the immediate flight control application of corrective action,

applying forward cyclic to regain airspeed, applying right pedal not left as necessary to maintain

rotor RPM, and reducing the collective (thus reducing the high-power demand on the tail rotor),

is the key to a safe recovery. Pilots should always set themselves up when conducting any

maneuver to have enough height and space available to recover in the event they encounter an

aerodynamic situation such as LTE.

Understanding the aerodynamic phenomenon of LTE is by far the most important factor in

preventing an LTE-related accident, and maintaining the ability and option either to go around if

making an approach or pull out of a maneuver safely and re-plan, is always the safest option.

Having the ability to fly away from a situation and re-think the possible options should always be

part of a pilot's planning process in all phases of flight. Unfortunately, there have been many

pilots who have idled a good engine and fully functioning tail rotor disk and autorotated a

perfectly airworthy helicopter to the crash site because they misunderstood or misperceived both

the limitations of the helicopter and the aerodynamic situation.



Figure 9-6 Effects of Wind Direction on Directional Control

4. Weathercock Stability (120 - 240 degrees)

In this region, the helicopter attempts to weathervane, or weathercock, its nose into the relative

wind. Unless a resisting pedal input is made, the helicopter starts a slow, un-commanded turn

either to the right or left, depending upon the wind direction. If the pilot allows a right yaw rate

to develop and the tail of the helicopter moves into this region, the yaw rate can accelerate

rapidly. In order to avoid the onset of LTE in this downwind condition, it is imperative to

maintain positive control of the yaw rate and devote full attention to flying the helicopter.

5. Tail Rotor VRS (210 - 330 degrees)

Winds within this region cause a tail rotor vortex ring state to develop. The result is a non-

uniform, unsteady flow into the tail rotor. The vortex ring state causes tail rotor thrust variations,

which result in yaw deviations. The net effect of the unsteady flow is an oscillation of tail rotor

thrust. Rapid and continuous pedal movements are necessary to compensate for the rapid

changes in tail rotor thrust when hovering in a left crosswind. Maintaining a precise heading in

this region is difficult, but this characteristic presents no significant problem unless corrective

action is delayed. However, high pedal workload, lack of concentration, and over controlling

can lead to LTE.

When the tail rotor thrust being generated is less than the thrust required, the helicopter yaws to

the right. When hovering in left crosswinds, concentrate on smooth pedal coordination and do

not allow an un-commanded right yaw to develop. If a right yaw rate is allowed to build, the

helicopter can rotate into the wind azimuth region where weathercock stability then accelerates

the right turn rate. Pilot workload during a tail rotor vortex ring state is high. Do not allow a

right yaw rate to increase.



6. Main Rotor Disk Vortex (285 - 315 degrees)

Winds at velocities of 10–30 knots from the left front cause the main rotor vortex to be blown

into the tail rotor by the relative wind. This main rotor disk vortex causes the tail rotor to operate

in an extremely turbulent environment. During a right turn, the tail rotor experiences a reduction

of thrust as it comes into the area of the main rotor disk vortex. The reduction in tail rotor thrust

comes from the airflow changes experienced at the tail rotor as the main rotor disk vortex moves

across the tail rotor disk.

The effect of the main rotor disk vortex initially increases the AOA of the tail rotor blades, thus

increasing tail rotor thrust. The increase in the AOA requires that right pedal pressure be added

to reduce tail rotor thrust in order to maintain the same rate of turn. As the main rotor vortex

passes the tail rotor, the tail rotor AOA is reduced. The reduction in the AOA causes a reduction

in thrust and right yaw acceleration begins. This acceleration can be surprising, since previously

adding right pedal to maintain the right turn rate. This thrust reduction occurs suddenly, and if

uncorrected, develops into an uncontrollable rapid rotation about the mast. When operating

within this region, be aware that the reduction in tail rotor thrust can happen quite suddenly, and

be prepared to react quickly to counter this reduction with additional left pedal input.

7. AOA Reduction (060 - 120 degrees)

In a right crosswind, the relative wind shifts toward a tail rotor blades’ chord line because of

effectively increased induced velocity. The shifted relative wind impacts at a lower AOA, which

develops lower lift and results in less thrust. The pilot will automatically compensate by adding

more left pedal, but in some cases can reach pedal travel limits before adequate thrust can be

generated.



Figure 9-7 LTE in a Right Crosswind

8. Loss of Translational Lift (All Azimuths)

The loss of translational lift results in increased power demand and additional anti-torque

requirements. If the loss of translational lift occurs when the aircraft is experiencing a right turn

rate, the right turn will be accelerated as power is increased unless corrective action is taken by

the pilot.

9. Reducing the Onset of LTE

To help reduce the onset of LTE, follow these steps:

a. Maintain maximum power-on rotor RPM. If the main rotor RPM is allowed to

decrease, the anti-torque thrust available is decreased proportionally.

b. Avoid tailwinds below airspeeds of 30 knots. If loss of translational lift occurs, it

results in an increased power demand and additional anti-torque pressures.

c. Avoid OGE operations and high-power demand situations below airspeeds of

30 knots at low altitudes.

d. Be especially aware of wind direction and velocity when hovering in winds of about

8–12 knots. A loss of translational lift results in an unexpected high power demand

and an increased anti-torque requirement.



e. Be aware that if a considerable amount of left pedal is being maintained, a sufficient

amount of left pedal may not be available to counteract an unanticipated right yaw.

f. Be alert to changing wind conditions, which may be experienced when flying along

ridge lines and around buildings.

g. Execute right turns slowly. This limits the effects of rotating inertia, and decreases

loading on the tail rotor to control yawing.

10. Recovery Technique

If a sudden unanticipated right yaw occurs, the following recovery technique should be

performed. Apply full left pedal. Simultaneously, apply forward cyclic control to increase

speed. If altitude permits, reduce power. As recovery is affected, adjust controls for normal

forward flight. A recovery path must always be planned, especially when terminating to an OGE

hover and executed immediately if an un-commanded yaw is evident.

Collective pitch reduction aids in arresting the yaw rate but may cause an excessive rate of

descent. Any large, rapid increase in collective to prevent ground or obstacle contact may

further increase the yaw rate and decrease rotor RPM. The decision to reduce collective must be

based on the pilot’s assessment of the altitude available for recovery.

If the rotation cannot be stopped and ground contact is imminent, an autorotation may be the best

course of action. Maintain full left pedal until the rotation stops, then adjust to maintain heading.

For more information on LTE, see Advisory Circular (AC) 90-95, Unanticipated Right Yaw in

Helicopters.

11. Tail Rotor Failure and Engine Failure

Loss of engine power. Should the aircraft lose power the aircraft will tend to yaw left. The yaw

inputs made prior to the engine failure compensate for a much greater torque than that which is

instantly delivered with a reduction in engine power. The tail rotor continues to provide thrust

whether it is powered by one engine in a single engine failure, or through windmilling in a

complete engine failure. Initial response to engine failures must include a right pedal input.

Tail rotor failure. Tail rotor failure, whether a control failure (stuck pedal) or a complete loss

of tail rotor thrust, can be a survivable event. With a control failure, most designs allow for the

tail rotor to operate at some intermediate setting. If the pilot chooses an appropriate speed,

balanced flight associated with that tail rotor setting can be attained.

A complete loss of tail rotor thrust requires more attention to airspeed. With increased airspeed

the main rotor operates more efficiently so it generates less torque. As velocity increases, both

the power required and anti-torque required decrease until the aircraft reaches its minimum

power required or "bucket" airspeed. After the bucket airspeed the power and anti-torque

required again increases up to VNE. The best airspeed to fly during a tail rotor failure would be

that requiring the least amount of anti-torque. An even better option is to fly in the flight regime



in which the engines produce no torque, that is, an autorotation. However, the impact of the

vertical stabilizer (fin) must be taken into account as indicated below.

A vertical stabilizer can help quite a bit in reducing the amount of tail rotor thrust required in

forward flight. Shaped like a wing, a vertical stabilizer provides lift (thrust) in the direction of

anti-torque. The higher the aircraft's velocity the more the vertical stabilizer will be contributing

to the anti-torque effort. Making a tail fin that could completely compensate for a lost tail rotor

would compromise sideward flight capability. The Apache and Blackhawk, however, were

designed to be able to fly straight in a controlled descent (at an appropriate airspeed) without the

tail rotor operating.

Figure 9-8 Fly Home Capability After Loss Of Tail Rotor Thrust



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GLOSSARY A-1

APPENDIX A

GLOSSARY

A100. GLOSSARY

Acceleration: The time rate of change of velocity.

Advancing blade: The rotor blade experiencing an increased relative wind because of airspeed.

Aerodynamics: 1The science that treats the motion of air and other gaseous fluids and the forces

acting on bodies when the bodies move through such fluids or when such fluids move against or

around the bodies. 2aThe actions and forces resulting from the movement or flow of gaseous

fluids against or around bodies. 2bThe properties of a body or bodies with respect to these

actions or forces. 3The application of the principles of gaseous fluid flows and their actions

against and around bodies to the design and construction of bodies intended to move through

such fluids.

Aerodynamic center (AC): Point along the chord line about which changes in AOA do not

result in a change of moment.

Aerodynamic force: The vector summation of lift and drag vectors depicted on the blade

element diagram.

Aerodynamic twist: The twist of an airfoil having different absolute angles of incidence at

different span-wise stations.

Air density/Atmospheric density: Mass of air per unit volume (D = M/V). It is the single most

important atmospheric variable with regards to aircraft performance.

Airfoil: A structure designed to produce lift as it moves through the air.

Airfoil characteristics: 1Any aerodynamic quality peculiar to a particular airfoil, especially to

an airfoil section or profile, usually a specified AOA. Airfoil characteristics are expressed

variously as the coefficients of lift or drag, the pitching moment, the zero-lift angle, the lift-drag

ratio, and so on. 2A feature of any particular airfoil or airfoil section such as the actual or

relative amount of span, taper, or thickness.

Airfoil section: 1A section of an airfoil, especially a cross section, taken at right angles to the

span axis or some other specified axis of the airfoil. 2The form or shape of an airfoil section; an

airfoil profile or the area defined by the profile.

Airspeed: The speed of an aircraft in relation to the air through which it is passing. Typically in

terms of forward airspeed but can be sideways or rearward also.

APPENDIX A HELICOPTER AERODYNAMICS WORKBOOK

A-2 GLOSSARY

Angle-of-Attack (AOA): The angle at which a body, such as an airfoil or fuselage, or a system

of bodies, such as a helicopter rotor, meets a flow. Usually expressed as the acute angle between

the chord line of an airfoil and the resultant relative wind.

Angle of climb: The angle between a horizontal plane and the flight path of a climbing aircraft.

Angle of incidence: Fixed airfoils (wings, horizontal and vertical fins, stabilizers): the acute

angle between the chord line of the airfoil and a selected reference plane, usually the longitudinal

axis of the aircraft. Rotating airfoils (helicopters’ main and tail rotors, propellers): the acute

angle between the chord line of the airfoil and the tip path plane. Angle of incidence is normally

called pitch angle for main rotor, tail rotor, and propeller blades.

Angular acceleration: A simultaneous change in both speed and direction of movement. An

example of this is an airplane in a spin.

Anti-autorotative force: In autorotational flight, the decelerating horizontal component of the

aerodynamic force along the driven and no-lift regions.

Anti-torque device: A method used to counteract torque reaction, for example a tail rotor,

Fenestron, or NOTAR to name a few.

Articulated rotor system: A rotor system in which the hub is mounted rigidly to the mast and

the individual blades are mounted on hinge pins, allowing them to flap up and down and move

forward and backward (lead and lag). Individual blades are allowed to feather by rotating about

the blade grip retainer bearing.

Aspect Ratio: Length of a blade divided by its width.

Attitude: The position of a body as determined by the inclination of the axes to some frame of

reference. If not otherwise specified, this frame of reference is fixed to the earth (horizon).

Autorotation: Descending flight of a helicopter without engine power where the air

approaching from below the rotor disk (upward induced flow) keeps the rotor blades turning at

an operational speed. May be divided into four distinct phases: entry, steady state descent, flare

and touchdown.

Axis: 1A line passing through a body about which the body rotates or may be assumed to rotate.

Any arbitrary line of reference such as a line about which the parts of a body or system are

symmetrically distributed. A line along which a force is directed; for example, an axis of thrust. 2Specifically, any one of a set or system of mutually perpendicular reference axes—usually

intersecting at the CG of an aircraft, rocket projectile, or the like—about which the motions,

moments, and forces of roll (longitudinal), pitch (lateral), and yaw (vertical) are measured.

Balancing tab: A moveable tab linked to the trailing edge of a control surface. When the

control surface is deflected the tab is deflected in an opposite direction, creating a force which

aids in moving the larger surface. Sometimes called a servo tab.

HELICOPTER AERODYNAMICS WORKBOOK APPENDIX A

GLOSSARY A-3

Blade element theory: Utilizes graphically depicted representation of the airflow and

aerodynamic forces applied to a selected airfoil section. Gives a more accurate representation of

rotor performance than does Momentum Theory. It also details the movement of individual

blades around the disk.

Blowback: The pitch-up tendency as the aircraft accelerates due to the flapping which

compensates for dissymmetry of lift. The separation of the virtual axis from the control axis.

Boundary-layer control: The control of the flow in the boundary layer about a body, or of the

region of flow near the surface of the body, to reduce or eliminate undesirable aerodynamic

effects and hence to improve performance.

Camber: The curvature of the surfaces of an airfoil or airfoil section from leading edge to

trailing edge.

Camber Line: Line equidistant from the upper and lower surface of the airfoil; same as chord

line for a symmetrical airfoil.

Center of gravity (CG): The balancing point for a body, generally expressed along the

longitudinal or lateral axis.

Center of pressure: Point along chord line about which all aerodynamic forces (distributed lift

along upper and lower surfaces) are acting.

Center-of-pressure travel: The movement of the center of pressure of an airfoil along the

chord with changing AOA; the amount of this movement is expressed in percentages of the

chord length from the leading edge.

Centrifugal force: The outward force created by the rotation of the main rotor and opposed by

centripetal force. The large centrifugal force is what allows the weight of the helicopter to be

distributed across otherwise flexible rotor blades. Centrifugal force is proportional to the square

of Nr and increases dynamic blade rigidity.

Centripetal force: The accelerative force acting on a body moving in a curved path. It is the

component of force that is directed toward the center of curvature or axis of rotation. Centripetal

force causes a change in the direction of the linear velocity vector of a body in motion, resulting

in an acceleration of the body. Centripetal force is the out-of-balance force that causes an

aircraft to turn. It is the horizontal component of lift that is directed toward the center of the

turn.

Chord: The distance between the leading and trailing edges of an airfoil along the chord line.

Chord line: A straight line intersecting the leading and trailing edges of an airfoil.

Coefficient of drag (CD): A dimensionless number indicating the inefficiency of an airfoil

which is determined by AOA and airfoil design. It is derived from wind tunnel testing.


A-4 GLOSSARY

Coefficient of lift (CL): A dimensionless number indicating the efficiency of the airfoil which is

determined by AOA and airfoil design. It is derived from wind tunnel testing.

Collective feathering: The equal and simultaneous mechanical change of blade pitch (the angle

of incidence) of all rotor blades in a rotor system.

Compressibility: At high forward airspeeds, the advancing rotor blade creates large pressure

changes, which result in significant air density changes. As the blade’s velocity approaches the

speed of sound, the blade becomes less efficient because of a nose-down pitching moment and a

significant increase in drag.

Compressible flow: Flow at speeds high enough that density changes in the fluid can no longer

be neglected.

Coning: The upward displacement of the main rotor blades due to increased lift and balanced

somewhat by centrifugal force.

Coning angle: The angle between the tip path plane and the main rotor blades.

Control surface: A movable airfoil designed to be rotated or otherwise moved to change the

speed or direction of an aircraft.

Critical Mach number: The free-stream Mach number at which a local Mach number of 1.0 is

attained at any point on the body under consideration.

Cyclic feathering: The mechanical change of blade pitch (the angle of incidence), of individual

rotor blades independently of the other blades in the system.

Density altitude (DA): PA corrected for temperature and humidity; or, the altitude in the

standard atmosphere corresponding to a particular value of air density. The denser an air mass

(cold, dry air), the lower the corresponding value corrected to a standard atmosphere will be

(High density = Low DA). The opposite is also true. Additionally, DA increases as temperature

and/or relative humidity increases. Therefore DA is inversely proportional to atmospheric

density and directly proportional to temperature and relative humidity.

Disk Area: The area of the circle inscribed by the tip path plane with the rotors turning. The

coning angle of the blades changes the disk area.

Disk loading: The weight (thrust) of the helicopter divided by the rotor disk area (lb./sq.in).

Dissymmetry of lift: In forward flight the advancing blade experiences an increase in linear

flow. The increased linear flow increases the lift on the advancing blade. Likewise, the

retreating blade sees a decrease in linear flow and therefore a decrease in lift. Compensated for

primarily by flapping.

Downwash: The induced downward flow of air resulting from the passage of an airfoil

(induced flow).


GLOSSARY A-5

Downwash angle: The angle, measured in a plane parallel to the plane of symmetry of an

aircraft, between the direction of downwash and the direction of the undisturbed airstream. This

angle is positive when the deflected stream is downward. (See Up-wash angle.)

Drag: The aerodynamic force in a direction opposite that of flight and caused by the resistance

to movement brought to bear on an aircraft by the atmosphere through which it passes.

Droop snoot airfoil: nonsymmetrical airfoil design used by the TH-57. The droop snoot design

incorporates a symmetrical blade design with a nonsymmetrical nose. A droop snoot design

provides good stall characteristics at high angles of attack and produces very low pitching

moments.

Dynamic pressure: The pressure of a fluid resulting from its motion; it is equal to one-half the

fluid density times the fluid velocity squared (q = 1/2V2). In incompressible flow, dynamic

pressure is the difference between total pressure and static pressure.

Dynamic rollover: The lateral rolling of the helicopter onto its side due to exceeding the critical

rollover angle for a critical roll rate, regardless of cyclic corrections. For dynamic rollover to

occur the helicopter must have a ground pivot point.

Dynamic stability: The property that causes a body, such as an aircraft or a rocket, to dampen

the oscillations set up by restoring moments and to return gradually to its original state when

disturbed from the original state of steady flight or motion.

Effective translational lift: The pronounced increased in translational lift during transition to

forward flight (approximately 13-24 knots) due to the rotor disk experiencing a significantly

decreased induced airflow.

Empennage: The assembly of stabilizing and control surfaces at the tail of an aircraft.

Endurance: The time an aircraft can continue flying under given conditions without refueling.

Equivalent airspeed: Calibrated airspeed of an aircraft corrected for adiabatic compressible

flow for the particular altitude. Equivalent airspeed is equal to calibrated airspeed in standard

atmosphere at sea level.

Feathering: A mechanical change in the angle of incidence, or pitch, of an airfoil segment.

Fin: A fixed airfoil that aids directional stability.

Flapping: Vertical blade movement, normally about a central hinge pin, which allows the rotor

disk to tilt and helps compensate for dissymmetry of lift.

Flight path: The line connecting the continuous positions occupied or to be occupied by an

aircraft as it moves with reference to the vertical or horizontal planes.


A-6 GLOSSARY

Flow separation/boundary layer separation: The breakaway of flow from a surface; the

condition of a flow separated from the surface of a body and no longer following its contours.

Fuselage: The body to which the wings, landing gear, and tail are attached.

Geometric imbalance: Occurs when the radius of the center of mass for a single rotor blade

changes due to excessive flapping and is no longer equidistant from the center of rotation relative

to the individual centers of mass for the other rotor blade(s). This phenomenon may lead to

excessive hunting. If the center of mass for an individual rotor blade shifts towards the center of

rotation, that blade has a tendency to lead; likewise, if the center of mass shifts away from the

center of rotation, the blades will have a tendency to lag. Ground resonance may result if this

excessive flapping creates excessive hunting oscillations.

Geometric twist: An engineered design of the rotor blade span-wise, that incorporates a twist

beginning with an increased angle of incidence at the root of the rotor blade which decreases

from the root to the tip. Geometric twist helps to distribute lift more equally across the rotor

blade.

Gravity: An attraction of two objects for each other that depends on their mass and the distance

between them.

Gross weight: The total weight of an aircraft and its contents.

Ground effect: The increased efficiency (decreasing power requirement) of the rotor system of

the helicopter beginning at approximately one rotor diameter above the surface and increasing as

the helicopter approaches the ground. The aerodynamic effect can be largely attributed to the

reduction of the velocity of the induced flow because the ground interrupts the airflow beneath

the helicopter. Additionally, the ground interrupts the formation of tip vortices, reducing their

contribution to induced flow. The decrease in induced flow increases AOA, providing an

increase in lift with a reduction in blade pitch setting/power setting.

Ground resonance: Normally associated with the fully articulated rotor system and an

inoperative blade dampener, ground resonance is a destructive oscillation caused when the

helicopter is in contact with the ground and one or more rotor blades are displaced due to a gust

of wind, sudden control movement, or a hard landing. When this occurs, the CG of the rotor

system spirals violently outward. (See Geometric imbalance)

Ground vortex: During a normal transition to forward flight, the helicopter’s downwash creates

a vortex in front of the path of flight. As the helicopter accelerates, the aircraft flies through the

vortex. This serves to increase the induced flow causing an increase in the power required.

Gyroscopic precession: A phenomenon in rotating systems that results in all forces applied

perpendicular to the plane of rotation being manifested 90° later from the point of force in the

direction of rotation.

Horsepower: A unit of power equal to the power necessary to raise 550 pounds one foot in one

second. Thus a 1000-horsepower engine develops 1000 times 550 foot-pounds of work per


GLOSSARY A-7

second. It is common to represent this power in terms of minutes instead of seconds. Thus,

equations routinely have conversion factors of 33,000.

Induced drag: The horizontal component of lift (parallel to the tip path plane) attributed to a

downward induced velocity.

Induced flow (V-ind): Vertical/axial component of relative wind. Generally, in powered flight

the induced velocity is downward and in non-powered flight the induced velocity is upward

through the rotor disk. Also known as induced airflow.

In-plane drag: The summation of all decelerating forces in the plane of rotation (induced drag +

horizontal component of profile drag).

Kinetic energy: The energy of a system because of motion.

Lag: In a rotating system, this is the occurrence of a momentary decrease in the rotational

velocity, normally about a vertical hinge pin in an articulated system.

Laminar flow: A smooth flow in which no cross flow of fluid particles occurs, hence a flow

conceived as made up of layers.

Laminar separation: The separation of a laminar-flow boundary layer from a body.

Lateral axis: An axis going from side to side of an aircraft, rocket, missile, and so on. It is

usually the side-to-side body axis passing through the CG. The axis about which pitching action

occurs. Sometimes called a Transverse axis.

Lateral stability: The tendency of a body, such as an aircraft, to resist rolling or, sometimes,

lateral displacement; the tendency of an aircraft to remain wings-level, either in flight or at rest.

Lead: Opposite of lag, or, a momentary increase in the rotational velocity in a rotating system.

Leading edge: The forward edge of an airfoil, blade, and the like. The edge which normally

meets the air or fluid first.

Lift: The component of the total aerodynamic force (thrust on a blade element), which is

perpendicular to the relative wind.

Lift component: A force acting on an airfoil perpendicular to the direction of its motion through

the air.

Lift-drag ratio: The ratio of lift to induced drag, obtained by dividing the lift by the induced

drag or the coefficient of lift by the coefficient of drag.

Linear flow: Horizontal/lateral component of resultant relative wind in a rotating system, the

V-rotational flow +/- the V-translational, adjusted for any existing wind condition.


A-8 GLOSSARY

Load: 1The forces acting on a structure. These may be static (as with gravity), dynamic (as with

centrifugal force), or a combination of static and dynamic. 2Used to describe an aircraft’s cargo.

Load factor: the sum of the loads on a structure, including the static and dynamic loads;

expressed in units of G.

Longitudinal acceleration: Acceleration substantially along the longitudinal axis of an aircraft,

a rocket, or the like.

Longitudinal axis: A straight line through the CG of an aircraft fore and aft in the plane of

symmetry.

Mach number: The ratio of the velocity of a body to that of sound in the surrounding medium.

Thus a Mach number of 1.0 indicates a speed equal to the speed of sound; 0.5, a speed one-half

the speed of sound; 5.0, a speed five times the speed of sound, and so on.

Mach wave: 1A shock wave theoretically occurring along a common line of intersection of all

the pressure disturbances emanating from an infinitesimally small particle moving at a

supersonic speed through a fluid medium; such a wave is considered to exert no changes in the

condition of the fluid it is passing through. The concept of the Mach wave is used in defining

and studying the realm of certain disturbances in a supersonic field of flow. 2A very weak shock

wave appearing, for example, at the nose of a very sharp body where the fluid undergoes no

substantial change in direction.

Maneuver: Any planned motion of an aircraft in the air or on the ground.

Maneuverability: The ease with which an aircraft will move out of its equilibrium position.

Maneuverability and stability are opposites.

Maximum endurance airspeed: The lowest point on the power required curve where the ratio

of lift versus drag is maximized (also called bucket airspeed).

Maximum range airspeed: The point where a line drawn from the origin (corrected for winds)

is tangent to the power required curve.

Maximum rate of climb airspeed: The lowest point on the power required curve. Ratio of lift

versus the drag is maximized thereby allowing for the greatest power excess. (Also referred to

as best rate of climb)

Mean aerodynamic chord: The chord of an imaginary rectangular airfoil that would have

pitching moments throughout the flight range the same as those of an actual airfoil or

combination of airfoils under consideration, calculated to make equations of aerodynamic forces

applicable.

Mean camber line: A line drawn halfway between the upper and lower surfaces of an airfoil.

The curvature of the mean camber line in relation to the chord line is very important in

determining the aerodynamic characteristics of an airfoil section. The maximum camber

(displacement of the mean line from the chord) and the location of the maximum camber help to


GLOSSARY A-9

define the shape of the mean camber line. These quantities are expressed as fractions or a

percent of the basic chord length. A typical low-speed airfoil may have a maximum camber of

4% located 40% aft of the leading edge. On symmetrical airfoils, the mean camber line and the

chord line are the same.

Mechanical axis: The extension of the centerline of the rotor mast (the actual axis of the rotor

head).

Momentum theory: Theory that helps explain rotary wing lift production, primarily based on

Isaac Newton’s three Laws of Motion. The action of accelerating a mass of air downward

produces a reaction that lifts the helicopter. Momentum theory is most applicable in hovering

and forward flight.

Neutral stability: The stability of a body such that after it is disturbed, it tends neither to return

to its original state nor to move further from it; that is, its motions or oscillations neither increase

nor decrease in magnitude.

Newton’s Laws of Motion:

1. Newton’s First Law (The Law of Equilibrium) “A body at rest tends to remain at rest and a

body in motion tends to remain in motion in a straight line at a constant velocity unless acted

upon by some unbalanced force.”

2. Newton’s Second Law (The Law of Acceleration) “The acceleration (a) of a body is

directly proportional to the force (F) exerted on the body, is inversely proportional to the mass

(m) of the body, and is in the same direction as the force.” "F = ma"

3. Newton’s Third Law (The Law of Interaction)” “For every action, there is an equal and

opposite reaction.”

Nonsymmetrical airfoil: An airfoil with a different shape or size above and below the chord

line.

Parasite drag: Drag incurred from components of an aircraft not contributing to lift.

Pendulum effect: Un-commanded nose-up tendency during deceleration that occurs in response

to an increase in collective pitch before mechanical and virtual axes are realigned. Compensated

for by pilot-induced feathering through forward cyclic.

Phase lag: When a rotating system in resonance receives a periodic excitation force sympathetic

with the natural frequency of the system, the response to the applied force is a maximum

displacement up to 90° after the force is applied. A phenomenon of the rotor system analogous

to gyroscopic precession which occurs as a result of a continuous excitation force.

Pitch angle: Angle between the chord line and the tip path plane. (See also Angle of incidence).

Pitching moment: A moment about a lateral axis of an aircraft, rocket, airfoil, and so on. This

moment is positive when it tends to increase the AOA or to nose the body upward.


A-10 GLOSSARY

Positive G: The foot ward inertial force produced by a headward acceleration. The force occurs

in a gravitational field or during an acceleration when the human body is so positioned that the

force of inertia acts on it in a head-to-foot direction.

Positive lift: Lift acting in an upward direction.

Potential energy: The energy of a system derived from position.

Power: The rate of doing work; often expressed in units of horsepower.

Power available (PA or Pavail): The amount of power an engine is capable of producing for given

conditions. As DA increases, engine power available decreases.

Power excess/Excess power: Ratio of power available to the power required. If the ratio is less

than 1 then power required exceeds the power available.

Power required (PR or Preq): The amount of power necessary to turn the rotor system at a

constant speed. As the DA increases, the pitch angle of the rotor blades must increase to

generate the same amount of lift. This creates more drag forces on the rotor system and therefore

more power is required to maintain a constant rotor speed.

Power required exceeds power available (PR > PA): An un-commanded rate of descent and/or

loss of rotor RPM caused by the power required exceeding the power available (also called

settling with power). Conditions that contribute to higher power required are high gross weights,

high G-loading, rapid maneuvering, high-density altitudes, loss of ground effect, and loss of

translational lift. High DA also contributes to loss of engine power available.

Power Settling: A term often used interchangeably with “settling with power” by different

services and texts. See “Vortex Ring State” and “power required exceeds power available” for

preferred terminology.

Preconing: The engineered design used to reduce stress associated with flexing on the root of

the rotor blades, the yoke, and the blade grips.

Pressure altitude (PA): The altitude of a given pressure in the standard atmosphere. See

Standard atmosphere. As pressure increases, density increases and DA decreases.

Pressure gradient: A change in the pressure of a gas or fluid per unit of distance.

Pro-autorotative force: In unpowered flight, the accelerating horizontal component of the total

aerodynamic force vector in the region where it is tilted forward of vertical/axial (driving

region).

Profile drag: Result of air friction acting on the blade element (parallel to the relative wind).

NOTE:

In a hover, profile drag accounts for 15-45 percent of the total

power consumption.


GLOSSARY A-11

Rate of climb: The rate at which an aircraft gains altitude; that is, the vertical component of its

airspeed in climbing.

Rate of descent: The rate at which an aircraft descends; that is, the vertical component of its

airspeed in descending; the rate at which a parachute and its burden descend.

Relative velocity: Velocity of the resultant relative wind.

Relative wind resultant: The vector resultant of the linear velocity + induced velocity as

depicted on the blade element diagram.

Retreating blade: The rotor blade experiencing a decreased relative wind because of airspeed.

Retreating blade stall: Aggravated case of dissymmetry of lift which results in the aircraft

pitching up and rolling left. As airspeed increases the retreating blade’s linear flow is reduced,

the blade flaps down, decreasing induced flow and increasing AOA. Eventually, as airspeed

increases further, the blade will exceed the critical AOA and will stall. With current blade

designs, a helicopter’s forward airspeed is primarily limited by retreating blade stall.

Reynolds number: The product of a typical length and the fluid speed divided by the kinematic

viscosity of the fluid. It expresses the ratio of the internal forces to the viscous forces.

Rigid rotor system: Sometimes referred to as “hingeless” since the rotor blades are fixed

rigidly to the hub without mechanical hinges for flapping, lead and lag (hunting), and on some

systems pitch change (feathering). Flapping and hunting occur through the flexing and bending

of the composite hub or “flextures.” Some systems also allow for pitch change through the

twisting of the materials rather than a pitch-change hub.

Roll: Movement around the longitudinal axis of an aircraft.

Rotational velocity (V-rot): The component of the relative wind produced by rotation of the

rotor blades i.e., the velocity of airflow across the airfoil due to its rotation about the mechanical

axis.

Rotor disk: Area of the circle inscribed in the tip path plane.

Rotor system: General term referring primarily to the design that holds the rotor blades to the

mast. The three general types of rotor systems are: fully-articulated, semi-rigid and rigid.

Semi-rigid rotor system: A rotor system in which the blades are connected to the mast by a

trunnion that allows blades to flap. Pitch change (feathering) is allowed at the hub about the

blade grip retainer bearing.

Separated flow: Flow over or about a body that has broken away from the surface of the body

and no longer follows its contours.


A-12 GLOSSARY

Settling with power: Also known as “Power Required Exceeds Power Available,” is a

hazardous helicopter flight condition in which the power required for a given maneuver or flight

regime is greater than the power available under the current ambient conditions. The terms

“settling with power” and “power settling” are used differently by Army and Navy helicopter

pilots, therefore the term “power required exceeds power available” is preferred. This should not

be confused with “Power Settling” which is more correctly called “Vortex Ring State.”

Sideslip: A movement of an aircraft such that the relative wind has a velocity component along

the lateral axis.

Skid: Rate of turn is greater than normal for a degree of bank established.

Slip: The rate of turn is less than normal for the degree of bank established.

Span: 1aThe dimension of an airfoil from end to end, from tip to tip, or from root to tip. 1bThe

dimension of an aircraft, measured between lateral extremities. 2The dimension of an airfoil

from tip to tip, measured in a straight line. Where ailerons or elevators extend beyond the tips of

the airfoil proper, their extension is included in the span. Sweeping an airfoil or giving it

dihedral decreases the span.

Speed: The rate at which an object moves in relation to time and distance.

Speed of sound: The speed at which sound travels in a given medium under specified

conditions.

Stabilator: A horizontal surface that pivots as a whole; it is distinct from the usual combination

of fixed and movable surfaces.

Stability: The property of an aircraft to maintain its attitude or to resist displacement and, if

displaced, to develop forces and moments tending to restore the original condition.

Stabilizer: A fixed or adjustable airfoil or vane that provides stability for an aircraft; that is, a

fin or more specifically the horizontal stabilizer on an aircraft.

Stall: 1aA condition in which a wing or other dynamically lifting body flies at an AOA greater

than that for maximum lift, resulting in a loss of lift and an increase of drag. 1bA loss of lift and

an increase of drag brought on by a shock wave; that is, a shock stall. 2The flight condition or

behavior of an aircraft flying at an angle greater than the angle of maximum lift; any of various

aircraft performances involving a stall.

Stall speed: The airspeed at which, under a given set of conditions, an aircraft will stall.

Stalling AOA: 1The minimum AOA of an airfoil or airfoil section or other dynamic lifting body

at which a stall occurs; that is, a critical AOA. 2The angle of maximum lift.

Standard atmosphere: A model of atmospheric conditions that vary with altitude above sea

level, namely: pressure, temperature, and density. The model was derived from global averages

and is used in performance.


GLOSSARY A-13

Standard lapse rate: In a thermodynamic system, the rate of heat loss of two degrees Celsius

per every 1000 feet due to an expansion of the atmosphere corresponding to an increase in

altitude. Also referred to as average or adiabatic lapse rate.

Static pressure: The atmospheric pressure of the air through which an aircraft is flying.

Sweepback: The backward slant from root to tip (or inboard end to outboard end) of an airfoil

or of the leading edge or other reference line of an airfoil. Sweepback usually refers to a design

in which both the leading and trailing edges of the airfoil have a backward slant.

Symmetrical airfoil: An airfoil with the same size and shape above and below the chord line.

Tab: A small auxiliary airfoil set into the trailing edge of an aircraft control surface (or

something set into or attached to another surface such as a rotor blade) and used for trim or to

move or assist in moving the larger surface.

Tail rotor: The anti-torque device of a single-rotor helicopter. Control of this rotor is through

the foot pedals.

Tandem rotor system: A main lifting rotor is used at each end of the helicopter. The rotor

systems rotate in opposite directions to counteract torque.

Taxi: 1The operation of an airplane or helicopter under its own power on the ground, except that

movement incident to actual takeoff and landing. 2The forward movement of a helicopter at a

hover is referred to as a hover taxi.

Thrust: Rotor thrust is the vector sum of forces produced in the rotor system.

Thrust axis: A line or axis through an aircraft, rocket, and so on along which the thrust acts; an

axis through the longitudinal center of a jet or rocket engine along which the thrust of the engine

acts; a center of thrust. For helicopters, the total rotor thrust acts perpendicular to the tip path

plane through the rotor head and is called virtual axis.

Tip path plane: The path inscribed by the tips of the main rotor blades as they rotate. The tip

path plane contains the rotor disk, and rotor thrust is perpendicular to the TPP.

Tip vortex: A vortex springing from the tip of a wing because of the flow of air around the tip

from the high-pressure region below the surface to the low-pressure region above it.

Torque: Mathematically, torque is a force times a distance. It causes the fuselage to react in

yaw due to the fact that the drive train turns the rotor.

Torque effect: In a counterclockwise rotating rotor system, due to the momentum of the

advancing rotor blade on the right side of the aircraft, there is an equal and opposite reaction

(torque) which causes the helicopter to rotate to the right. The tail rotor counteracts torque

effect. Remember Newton’s Third Law of Motion which states that every action has an equal

and opposite reaction.


A-14 GLOSSARY

Trailing edge: The rearmost edge of an airfoil.

Trailing vortex: A vortex that is shed from a wing or other lifting body and is trailing behind it,

especially such a vortex trailing from a wingtip or from the end of a bound vortex. It is

sometimes referred to as wake turbulence.

Translating tendency: Tendency for a helicopter to translate laterally due to tail rotor thrust.

Translational flight: Any horizontal movement of a helicopter with respect to the air.

Translational lift: The increased efficiency of the rotor system in the production of lift by

increasing the horizontal mass flow of air through the rotor disk, reducing the induced flow and

vortices. (See also Effective translational lift)

Translational velocity (V-trans): Airflow through a rotor system or across a blade element due

to movement of the aircraft. Added geometrically to v-rotational on the advancing blade and

subtracted on the retreating blade.

Transverse flow effect: A non-uniform induced velocity flow pattern across the rotor disk that

produces a pronounced rolling tendency and lateral vibrations during transition through

approximately 10-20 knots.

Trim: The condition of a heavier-than-air aircraft in which it maintains a fixed attitude with

respect to the wind axes, with the moments about the aircraft axes being in equilibrium. The

word “trim” is often used with special reference to the balance of control forces.

Trim tab: A tab that is deflected to a position where it remains to keep the aircraft in the desired

trim. Adjustment of a trim tab on a rotor blade causes the blade to maintain a given track or

plane of motion.

True airspeed: Equivalent airspeed corrected for error that is due to air density (altitude and

temperature).

Turbulence: An agitated condition of the air or other fluids; a disordered, irregular, mixing

motion of a fluid or fluid flow such as that about a body in motion through the air.

Turbulent boundary layer: A boundary layer characterized by random fluctuations of a

velocity and by pronounced layer mixing of the fluid.

Turbulent flow: A flow characterized by turbulence; that is, an irregular, eddying, fluctuating

flow; a flow in which the velocity of a given point varies erratically in magnitude and direction

with time.

Underslinging: Attachment of the rotor head occurs with a pivot point above the blade grips

and centered midway between the opposing blade centers of gravity. Semi-rigid rotor head

design which compensates for geometric imbalance by keeping the individual centers of mass for


GLOSSARY A-15

each rotor blade equidistant from the center of rotation. Allows for flapping, but geometric

design minimizes hunting.

Uniform flow: An idealized flow in which the streamlines are parallel and the velocity is

constant throughout.

Unsteady flow: A flow whose velocity components vary with time at any point in the fluid.

Unsteady flow is of fixed pattern if the velocity at any point changes in magnitude but not

direction and of variable pattern if the velocity at any point changes in direction.

Up-wash: A flow deflected upward by a wing, rotor, rotor blade, and so on.

Up-wash angle: A negative downwash angle; that is, the acute angle, measured in a plane

parallel to the plane of symmetry of an aircraft, between the direction of up-wash and the

direction of the undisturbed airstream.

Useful load: The difference, in pounds, between the empty weight and maximum authorized

gross weight of an aircraft.

V-rotational: See “Rotational velocity.”

V-translational: See “Translational velocity.”

Vector: A quantity having both magnitude and direction. Also a graphic illustration of such a

quantity.

Velocity: 1Speed. 2A vector quantity that includes both magnitude (speed) and direction relative

to a given frame of reference. 3Time rate of motion in a given direction.

Venturi: A converging-diverging passage for fluid that increases the fluid velocity and lowers

its static pressure; a venturi tube.

Vertical axis: An axis passing through an aircraft from top to bottom and usually passing

through the CG. The axis about which yaw occurs. Also called a Normal axis.

Vertical stabilizer: A vertical fin mounted approximately parallel to the longitudinal axis of an

aircraft to which a rudder may be attached. The vertical stabilizer aids in directional stability.

Also called a vertical fin.

Virtual axis: The axis of rotation perpendicular to the tip path plane, as opposed to the

mechanical axis. As the rotor disk tilts with control inputs, the virtual axis tilts and remains

perpendicular to the plane of rotation. Rotor thrust acts through the virtual axis.

Vortex Ring State (VRS): Settling of the helicopter into its own downwash. During VRS,

airflow is downward over the outer portion of the rotor disk and upward both in an area

expanding outward from the hub as well as the area outside the tip path plane. This rapidly

decaying phenomenon may result in zero net lift. The prescribed limits to avoid entry into VRS


A-16 GLOSSARY

for the TH-57 are: avoid descent rates in excess 800 ft. /min at airspeeds less than 40 KIAS, and

avoid descent gradients greater than 45°. VRS has also been called “power settling,” a term

commonly confused with the term “settling with power.”

Wake Turbulence: (See Trailing vortex.)

Weathervane: The tendency of an aircraft on the ground to face into the wind.

Weight: A measure of the mass of an object under the acceleration of gravity.

Work: A force exerted over a given distance.

Yaw: A movement about the vertical axis.

Zero AOA: The position of an airfoil, fuselage, or other body when no AOA exists between two

specified or understood reference lines.

Zero-lift AOA: The geometric AOA at which no lift is created. Often called the angle of zero

lift or the zero-lift angle.

REFERENCES AND READING LIST B-1

APPENDIX B

REFERENCES AND READING LIST

1. Rotary Wing Aerodynamics for Naval Aviators, School of Aviation Safety, Department of

the Navy, May 2004.

2. Fundamentals of Aerodynamics (NAVAVSCOLSCOM-SG-111), Naval Aviation Schools

Command, Pensacola, FL, April 2008.

3. Fundamentals of Flight, FM 1-203, Headquarters, Department of the Army,

October 1988.

4. Rotary Wing Flight, FM 1-51, Headquarters, Department of the Army, 16 April 1979.

5. Meteorology for Army Aviators, FM 1-230, Headquarters, Department of the

Army, September 1982.

6. Anderson, John D, Fundamentals of Aerodynamics - 2nd Ed., McGraw-Hill

Inc., U.S.A. 1991, 1984.

7. Anderson, John D. Jr. A History of Aerodynamics - Cambridge University Press,

Cambridge, United Kingdom, 1999.

8. Dole, Charles E., Flight Theory and Aerodynamics - A Practical Guide for

Operational Safety, John Wiley & Sons, Inc. 1981.

9. Hurt, Hugh. Aerodynamics for Naval Aviators (NAVWEPS 00-80T-80).

Washington, D.C.: U.S. Government Printing Office, 1960.

10. Johnson, Wayne, Helicopter Theory Dover Publications, Inc., New York New York, 1980.

11. Leishman, J. Gordon, Principles of Helicopter Aerodynamics, Cambridge University

Press, Cambridge, United Kingdom, 2000.

12. Montgomery, John R, Sikorsky Helicopter Flight Theory for Pilots and

Mechanics, Sikorsky Aircraft, Division of United Aircraft Corporation, U.S.A. 1964.

13. Prouty, R.W., Helicopter Aerodynamics, PJS Publications Inc., Peoria IL, 1985.

14. Prouty, R.W., Even More Helicopter Aerodynamics, PJS Publications Inc., Peoria IL,

1985.

15. Prouty, R.W., Helicopter Performance, Stability. And Control, Krieger

Publishing Company, Inc., Malabar, FL, 1995.

16. Rolls-Royce, The Jet Engine, Rolls-Royce Limited, 1969.

APPENDIX B HELICOPTER AERODYNAMICS WORKBOOK

B-2 REFERENCES AND READING LIST

17. Saunders, George H., Dynamics of Helicopter Flight, John Wiley & Sons, lnc., 1975.

18. Seddon, J., Basic Helicopter Aerodynamics, AIAA, 1990.

19. Spence, Charles F., editor, AIM/FAR 1998, McGraw-Hill, 1998.

20. Stepniewski, W. Z., Keys, C. N., Rotary-Wing Aerodynamics, Dover

Publications, Inc., 31 East 2nd Street. Mineola, N.Y. 11501

21. Treager, Irwin E., Aircraft Gas Turbine Engine Technology, McGraw-Hill, Inc.,

1970.

22. Wagtendonk, Walter J., Principles of Helicopter Flight, Aviation Supplies &

Academics, Inc., Newcastle, Washington, 1996.

23. Helicopter Flying Handbook, U.S. Department of Transportation, Federal Aviation

Administration, Flight Standards Service, FAA-H-8083-21B, 2019.

Documents

WORKBOOK - Naval Air Training Command