The main difference between helicopter and an airplane is the main source of lift. The airplane derives its lift from a fixed airfoil surface while the helicopter derives lift from a rotating airfoil called the rotor. Hence, the aircraft will be classified as either “fixed – wing” or “rotating wing.” The word “heli­copter” is derived from the Greek words meaning "helical wing” or "rotating wing."

Lift generation by a “rotating wing” enables the helicopter to accomplish its unique mission of hovering motionless in the air, taking off and landing in a confined or restricted area, and autorotating to a safe landing following a power failure. Lift generation by “rotating wing” is also responsible for some of the unusual problems the helicopter can encounter. Since the helicopter problems are due to par­ticular nature of the rotor aerodynamics, the basic flow conditions within the rotor must be considered in detail. For simplicity, the initial discussion will consider only the hover­ing rotor. Although the term hovering usually means remaining over a particular spot on the ground, it shall be considered here as flight at zero airspeed. This is necessary because the aerodynamic characteristics of the rotor depend on its motion with respect to the air and not the ground. Hovering in a 20 knot wind is aerodynamically equivalent to flying at an airspeed of 20 knots in a no-wind condition, and the characteristics will be identical in the two conditions.

The first point to realize is that the rotor is subject to the same physical laws of aero­dynamics and motion that govern flight of the fixed-wing airplane. The manner in which the rotor is subject to these laws is much more complicated due to the complex flow con­ditions.

Rotor lift can be explained by either of two methods. The first method, utilizing simple momentum theory based on Newton’s Laws, merely states that lift results from the rotor accelerating a mass of air downward in the same way that the jet engine develops thrust by accelerating a mass of air out the tailpipe. The second method of viewing rotor lift concerns the pressure forces acting on the various sections of the blade from root to tip. The simple momentum theory is useful in determining only lift characteristics while the “blade element’’ theory gives drag as well as lift characteristics and is useful in giving a picture of the forces at work on the rotor. In the “blade element” theory, the blade is divided up into “blade elements” as shown in figure 6.15. The forces acting on each blade element are analyzed. Then the forces on all elements are summed up to give the characteristics of the whole rotor. The relative wind acting on each segment is the resultant of two velocity components: (1) the velocity due to the rotation of the blades about the hub and (2) the induced velocity, or downwash velocity caused by the rotor, the velocity due to rotation at a particular element is proportional to the rotor speed and the distance of the element from the rotor hub.

Thus, the velocity due to rotation varies line­arly from zero at the hub to a maximum at the tip. A typical blade section with the forces acting on it is shown in figure 6.15.

A summation of the forces acting perpen­dicular to the plane of rotation (tip path plane) will determine the rotor thrust (or lift) char­acteristics while summation of the moments resulting from forces acting in the plane of rotation will determine the rotor torque char­acteristics. As a result of this analysis, the rotor thrust (or lift) is found to be propor­tional to the air density, a nondimensional thrust coefficient, and the square of the tip speed, or linear speed of the tip of the blade. The thrust coefficient is a function of the aver­age blade section lift coefficient and the rotor solidity, which is the proportion of blade area to disc area. The lift coefficient is identical to that used in airplane aerodynamics while the solidity is analagous to the aspect ratio in air­plane aerodynamics. The rotor torque is found to be proportional to a nondimensional torque coefficient, the air density, the disc atea, the square of the tip speed, and the blade radius. The torque coefficient is dependent upon the average profile drag coefficient of the blades, the blade pitch angle, and the average lift coefficient of the blades. The torque can be thought to result from components of profile and induced drag forces acting on the blades, similar to those on an airplane.

As in the airplane, there is one angle of attack or blade pitch condition that will result in the most efficient operation. Unfortu­nately, the typical helicopter rotor operates at a near constant RPM and thus a constant true airspeed and cannot operate at this most efficient condition over a wide range of altitude and gross weight as the fixed-wing airplane. The airplane is able to maintain an efficient angle of attack at various altitudes and gross weights by flying at various airspeeds but the helicopter will operate with a near constant rotor velocity and vary blade angle to contend with variations in altitude and gross weight.

If the rotor could operate within a wide range of rotor speed, the efficiency and per­formance could be improved.

With the previous relationships established for the rotor in hovering flight, the effect of forward flight or rotor translation can be con­sidered. With forward flight, a third velocity component, that of the forward velocity of the helicopter, must be considered in determining the relative wind acting on each rotor blade element. Since the entire rotor moves with the helicopter, the velocity of air passing over each of the elements on the advanc’ng blade is increased by the forward speed of the heli­copter and the velocity of the air passing over each element of the retreating blade is de­creased by the same amount. This is shown in figure 6.16.

If the blade angles of attack on both advanc­ing and retreating blades remained the same as in hovering flight, the higher velocity on the advancing blade would cause a dissym­metry of lift and the helicopter would tend to roll to the left. It was this effect that created great difficulty during many early helicopter and autogiro projects. Juan De La Cierva was the first to realize what caused this effect and he solved the problem by mounting his auto­giro blades individually on flapping hinges, thus allowing a flapping action to automati­cally correct the dissymmetry of lift that re­sulted from forward flight. This is the method still used in an articulated rotor system today. The see-saw, or semi-rigid, rotor corrects the lift dissymmetry by rocking the entire hub and blades about a gimbal joint. By rocking the entire rotor system forward, the angle of at­tack on the advancing blade is reduced and the angle of attack on the retreating blade is increased. The rigid rotor must produce cyclic variation of the blade pitch mechanically as the blade rotates to eliminate the lift dissym­metry. Irrespective of the method used to correct the dissymmetry of lift, identical aero­dynamic characteristics result. Thus, what is said about rotor aerodynamics is equally valid for all types of rotor systems.

By analyzing the velocity components acting on the rotor blade sections from, the blade root to the tip on both advancing and retreat­ing blades, a large variation of blade section angle of attack is found. Figure 6.16 illus­trates a typical variation of the local blade angle of attack for various spanwise positions along the advancing and retreating blades of a rotor at high forward speed. There is a region of positive angles of attack resulting in positive lift over the entire advancing blade. Immediately next to the hub of the retreating blade there is an area of reversed flow where the velocity due to the forward motion of the helicopter is greater than the rearward velocity due to the blade rotation. The next area is a negative stall region where, although the flow is in the proper direction relative the blade, the angle of attack exceeds that for negative stall. Progressing out the retreating blade, the blade angle of attack becomes less negative, resulting in an area of negative lift. Then the blade angle becomes positive again, resulting in a positive lift region. The blade angle continues to increase until near the tip of the retreating blade the positive stall angle of attack is exceeded, resulting in stalling of the tip section. This wide variation in blade section angles of attack results in a large variation in blade section lift and drag coeffi­cients. The overall lift force on the left and right sides of the rotor disc are equalized by cyclically varying the blade pitch as explained previously, but the drag variation is not eliminated. This drag variation causes a shaking force on the rotor system and con­tributes to the vibration of the helicopter.

RETREATING BLADE STALL. Retreat­ing blade stall results whenever the angle of attack of the blade exceeds the stall angle of attack of the blade section. This condition occurs in high speed flight at the tip of the retreating blade since, in order to develop the same lift as the advancing blade, the retreating

blade must operate at a greater angle of attack. If the blade pitch is increased or the forward speed increased the stalled portion of the rotor disc becomes larger with the stall progressing in toward the hub from the tip of the retreating blade. When approximately 15 percent of the rotor disc is stalled, control of the heli­copter will be impossible. Flight tests have determined that control becomes marginal and the stall is considered severe when the outer one-quarter of the retreating blade is stalled. Retreating blade stall can be recognized by rotor roughness, erratic stick forces, a vibration and stick shake with a frequency determined by the number of blades and the rotor speed. Each of the blades of a three-bladed rotor will stall as it passes through the stall region and create a vibration with three beats per rotor revolution. Other evidence of retreating blade stall is partial or complete loss of control or a pitch-up tendency which can be uncontrollable if the stall is severe.

Conditions favorable for the occurrence of retreating blade stall are those conditions that result in high retreating blade angles of attack. Each of the following conditions results in a higher angle of attack on the retreating blade and may contribute to retreating blade stall:

1. High airspeed

2. Low rotor RPM—operation at low

rotor RPM necessitates the use of higher blade pitch to get a given thrust from the rotor, thus a higher angle of attack

3. High gross weight

4. High density altitude

5- Accelerated flight, high load factor

6. Flight through turbulent air or gusts—

sharp updrafts result in temporary increase in blade angle of attack

7. Excessive or abrupt control deflections

during maneuvers

Recovery from a stalled condition can be effected only by decreasing the blade angle of attack below the stall angle. This can be accomplished by one or a combination of the following items depending on severity of the stall:

1. Decrease collective pitch

2. Decrease airspeed

3. Increase rotor RPM

4. Decrease severity of accelerated ma­

neuver or control deflection If the stall is severe enough to result in pitch-up, forward cyclic to attempt to control pitch-up is ineffective and may aggravate the stall since forward cyclic results in an increase in blade angle of attack on the retreating blade. The helicopter will automatically recover from a severe stall since the airspeed is decreased in the nose high attitude but recovery can be assisted by gradual reduction in collective pitch, increasing RPM, and leveling the heli­copter with pedal and cyclic stick.

From the previous discussion, it is apparent that there is some degree of retreating blade stall even at moderate airspeeds. However, the helicopter is able to perform satisfactorily until a sufficiently large area of the rotor disc is stalled. Adequate warning of the impend­ing stall is present when the stall condition is approached slowly. There is inadequate warning of the stall only when the blade pitch or blade angle of attack is increased rapidly. Therefore, unintentional severe stall is most likely to occur during abrupt control motions or rapid accelerated maneuvers.

COMPRESSIBILITY EFFECTS. The highest relative velocities occur at the tip of the ad­vancing blade since the speed of the helicopter is added to the speed due to rotation at this point. When the Mach number of the tip section of the advancing blade exceeds the critical Mach number for the rotor blade sec­tion, compressibility effects result. The criti­cal Mach number is reduced by thick, highly cambered airfoils and critical Mach number decreases with increased lift coefficient. Most helicopter blades have symmetrical sections and therefore have relatively high critical Mach numbers at low lift coefficients. Since the principal effects of compressibility are the

large increase in drag and rearward shift of the airfoil aerodynamic center, compressibility ef­fects on the helicopter increase the power re­quired to maintain rotor RPM and cause rotor roughness, vibration, stick shake, and an un­desirable structural twisting of the blade.

Since compressibility effects become more severe at higher lift coefficients (higher blade angles of attack) and higher Mach numbers, the following operating conditions represent the most adverse conditions from the stand­point of compressibility:

1. High airspeed

2. High rotor RPM

3. High gross weight

4. High density altitude

5. Low temperature—the speed of sound

is proportional to the square root of the absolute temperature. Therefore, sonic velocity will be more easily obtained at low temperatures when the sonic speed is lower.

6. Turbulent air—sharp gusts momen­

tarily increase the blade angle of attack and thus lower the critical Mach number to the point where compressibility effects may be en­countered on the blade.

Compressibility effects will vanish by de­creasing the blade pitch. The similarities in the critical conditions for retreating blade stall and compressibility should be noticed but one basic difference must be appreciated— compressibility occurs at HIGH RPM while retreating blade stall occurs at LOW RPM. Recovery technique is identical for both with the exception of RPM control.

AUTOROTATION CHARACTERISTICS. One of the unique characteristics of helicopters is their ability to take part of the energy of the airstream to keep the rotor turning and glide down to a landing with no power. Consideration of the rotor during a vertical autorotation will provide an understanding of why the rotor continues to rotate without power. During autorotation, the flow of air is upward through the rotor disc and there is a vertical velocity component equal to the rate of descent of the helicopter. In addition, there is a velocity component due to rotation of the rotor. The vector sum of these two velocities is the relative wind for the blade element. The forces resulting from the relative wind on each particular blade section will provide the reason why the rotor will continue to operate without power. First, consider a blade element near the tip of the blade as illus­trated in figure 6.17- At this point there is a lift force acting perpendicular to the relative wind and a drag force acting parallel to the relative wind through the aerodynamic center. Since the rotation of the rotor is affected only by forces acting in the plane of rotation, the important forces are components of the lift and drag force in the plane of rotation. In this low angle of attack high speed tip section, the net in-plane force is a drag force which would tend to retard the rotor. Next, con­sider a blade section at about the half-span position as illustrated in figure 6.17- In this case, the same forces are present, but the in­plane component of lift force is greater than the drag force and this results in a net thrust or forward force in the plane of rotation which tends to drive the rotor.

During a steady autorotation, there is a balance of torque from the forces along the blade so that the RPM is maintained in equi­librium at some particular value. The region of the rotor disc where there is a net drag force on the blade is called the "propeller region” and the region of the rotor disc where there is a net in-plane thrust force is called the ‘‘autorotation region.” These regions are shown for vertical autorotation and forward speed (or normal) autorotation in figure 6.17. Forces acting on the rotor blades in forward flight autorotation are similar to those in vertical autorotation but the difference will consist mainly of shifts of the autorotation region to the left and the addition of reverse


flow and negative stall regions similar to the powered flight condition.

Autorotation is essentially a stable flight condition. If external disturbances cause the rotor to slow down, the autorotation region of the disc automatically expands to restore the rotor speed to the original equilibrium condi­tion. On the other hand, if an external disturbance causes the rotor to speed up, the propeller region automatically expands and tends to accelerate the rotor to the original equilibrium condition. Actually the stable autorotation condition will exist only when the autorotational speed is within certain limits. If the rotor speed is allowed to slow some excessive amount, then the rotor becomes unstable and the RPM will decrease even further unless the pilot immediately corrects the condition by proper control action.

In case of engine failure, the fixed-wing airplane will be glided at maximum lift-drag ratio to produce maximum glide distance. If minimum rate of descent is desired in power-off flight rather than maximum glide distance, the fixed-wing airplane will be flown at some lower airspeed. Actually, the minimum rate of descent will occur at minimum power required. The helicopter exhibits similar char­acteristics but ordinarily the best autorotation speed may be considered that speed that results in the minimum rate of descent rather than maximum glide distance. The aerodynamic condition of the rotor which produces mini­mum rate of descent is:

Maximum ratio of

(Mean blade lift coefficient)3^ Mean blade drag coefficient

It is this ratio which determines the auto­rotation rate of descent. Figure 6,18 illus­trates the variation of autorotation rate of descent with equivalent airspeed for a typical helicopter. Point A on this curve defines the point which produces autorotation with mini­mum rate of descent. Maximum glide distance during autorotation descent would be obtained at the flight condition which produces the greatest proportion between airspeed and rate of descent. Thus, a straight line from the origin tangent to the curve will define the point for maximum autorotative glide dis­tance. This corresponds to Point В of figure 6.18. If the helicopter is bying glided at the speed for maximum glide distance, a decrease in airspeed would reduce the rate of descent but the glide distance would decrease. If the helicopter is being glided at the speed for minimum rate of descent, the rate of descent (steady state) can not be reduced but the glide distance can be increased by increasing the glide speed to that for maximum distance. Weight and wind affect the glide character­istics of a helicopter the same way an airplane is affected. Ideally, the helicopter autorotates at a higher equivalent airspeed at higher gross weight or when autorotating into a headwind.

In addition to aerodynamic forces which act on the rotor during autorotation, inertia forces are also important. These effects are usually associated with the pilot’s response time be­cause the rate a pilot reacts to a power failure is quite critical. The time necessary to reduce collective pitch and enter autorotation be­comes critical if the rotor inertia characteristics are such as to allow the rotor to slow down to a dangerous level before the pilot can react. With power on, the blade pitch is relatively high and the engine supplies enough torque to overcome the drag of the blades. At the instant of power failure the blades are at a high pitch with high drag. If there is no engine torque to maintain the RPM, the rotor will decelerate depending on the rotor torque and rotor inertia. If the rotor has high rotational energy the rotor will lose RPM less rapidly, giving the pilot more time to reduce collective pitch and enter autorotation. If the rotor has low rotational energy, the rotor will lose RPM rapidly and the pilot may not be able to react quickly enough to prevent a serious loss of rotor RPM. Once the collective pitch is at

the low pitch limit, the rotor RPM can be in­creased only by a sacrifice in altitude or air­speed. If insufficient altitude is available to exchange for rotor speed, a hard landing is inevitable. Sufficient rotor rotational energy must be available to permit adding collective pitch to reduce the helicopter’s rate of descent before final ground contact.

In the case of most small helicopters, at least 300 feet of altitude is necessary for an average pilot to set up a steady autorotation and land the helicopter safely without damage. This minimum becomes 500 to 600 feet for the larger helicopters, and will be even greater for helicopters with increased disc loading. These characteristics are usually presented in the flight handbook in the form of a “dead man’s curve" which shows the combinations of air­speed and altitude above the terrain where a successful autorotative landing would be diffi­cult, if not impossible.

A typical “dead man’s curve’’ is shown in figure 6.18. The most critical combinations are due to low altitude and low airspeed illus­trated by area A of figure 6.18. Less critical conditions exist at higher airspeeds because of the greater energy available to set up a steady autorotation. The lower limit of area A is a finite altitude because the helicopter can be landed successfully if collective pitch is held rather than reduced. In this specific case there is not sufficient energy to reach a steady state autorotation. The maximum altitude at which this is possible is approximately ten feet on most helicopters.

Area В on the “dead man’s curve” of figure 6.18 is critical because of ground contact flight speed or rate of descent, which is based on the strength of the landing gear. The average pilot may have difficulty in successfully flaring the helicopter from a high speed flight con­dition without allowing the vail rotor to strike the ground or contacting the ground at an ex­cessive airspeed. A less critical zone is some­times shown on this curve to indicate that higher ground contact speeds can be permitted when the landing surface is smooth. In ad­dition, various stability and control character­istics of a helicopter may produce critical con­ditions in this area. The critical areas of the “dead man’s curve" should be avoided unless such operation is a specific mission require­ment.

POWER SETTLING. The term “power settling” has been used to describe a variety of flight conditions of the helicopter. True “power settling” occurs only when the heli­copter rotor is operating in a rotary flow condition called the “vortex ring state.” The flow through the rotor in the “vortex ring state" is upward near the center of the disc and downward in the outer portion, resulting in a condition of zero net thrust on the rotor. If the rotor thrust is zero, the helicopter is effectively free-falling and ex­tremely high rates of descent can result.

The downwash distribution within the rotor is shown in figure 6.19 for the conditions of normal hovering and power settling. Part A of figure 6.19 illustrates the typical down – wash distribution for hovering flight. If sufficient power were not available to hover at this condition, the helicopter would begin to settle at some rate of descent depending on the deficiency of power. This rate of descent would effectively decrease the downwash throughout the rotor and result in a redistri­bution of downwash similar to Part В of figure 6.19. At the outer portion of the rotor disc, the local induced downwash veloc­ity is greater than the rate of descent and downflow exists. At the center of the rotor disc, the rate of descent is greater than the local induced downwash velocity and the resultant flow is upward. This flow condition results in the rotary “vortex ring” state. By reference to the basic momentum theory it is apparent that the rotor will produce no thrust in this condition if the net mass flow of air through the rotor is zero. It is important to note that the main lifting part of the rotor is not stalled. The rotor roughness and loss of





control experienced during “power settling" results from the turbulent rotational flow on the blades and the unsteady shifting of the flow in and out span wise along the blade. There is an area of positive thrust in the outer portion of the rotor as a result of the mass of air accelerated downward and an area of negative thrust at the center of the rotor as a result of the mass of air flowing upward. The rotor is stalled only near the hub but no important effect is contributed because of the low local velocities.

Operation in the "vortex ring” state is a transient condition and the helicopter will seek equilibrium by descending. As the heli­copter descends, a greater upflow through the disc results until eventually the flow is entirely up through the rotor and the rotor enters auto­rotation where lower rates of descent can be achieved. Unfortunately, considerable alti­tude will be lost before the autorotative type of flow is achieved and a positive recovery technique must be applied to minimize the loss of altitude. "Power settling" can be recog­nized by rotor roughness, loss of control due to the turbulent rotational flow, and a very high rate of descent (as high as 3,000 fpm). It is most likely to be encountered inadvert­ently when attempting to hover when suf­ficient power is not available because of high gross weight or high density altitude.

Recovery from “power settling" can be ac­complished by getting the rotor out of the “vortex ring state." If the condition is en­countered with low power, rapid application of full power may increase the downwash suf­ficiently to get the rotor out of the condition. If the condition is encountered at high or maximum power or, if maximum power does not effect a recovery, increasing airspeed by diving will result in recovery with minimum loss of altitude. This type of recovery is most effective but adequate cyclic control must be available. If cyclic control has been lost, re­covery must be effected by reducing power and collective pitch and entering autorotation.

When normal autorotation has been estab­lished, a normal power recovery from the auto­rotation can be made. While such a recovery technique is effective, considerable altitude may be lost. Hence, diving out of the power settling condition provides the most favorable means of recovery.

Actually, real instances of true “power settling" are quite rare. A condition often described incorrectly as “power settling" is merely a high sink rate as a result of insufficient power to terminate an approach to landing. This situation frequently occurs during high gross weight or high density altitude operation. The flow conditions within the rotor are quite normal and there is merely insufficient power to reduce rate of descent and terminate an approach. Such a situation becomes more critical with a steep approach since the more rapid descent will require more power to terminate the approach.

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