Category AVIATORS

For the professional aviator, there are few documents which are as important as the air­plane flight handbook. The information and data contained in the various sections of the flight handbook provide the basis for safe and effective operation of the airplane.

Various sections of the flight handbook are devoted to the following subjects:

(1) Equipment and Systems. With the me­chanical complexity of the modern airplane, it is imperative that the pilot be familiar with every item of the aircraft. Only through exact knowledge of the equipment can the pilot properly operate the airplane and contend with malfunctions.

(2) Operating Procedures. Good procedures are mandatory to effect safe operation of the airplane and its equipment. The complexity of modern equipment dictates the use of special and exact procedures of operation and any haphazard or non-standard procedure is an

invitation for trouble of many sorts. The normal and emergency procedures applicable to each specific airplane will insure the proper operation of the equipment.

(3) Operating Limitations. The operation of the airplane and powerplant must be conducted within the established limitations. Failure to do so will invite failure or malfunction of the equipment and increase the operating cost or possibly cause an accident.

(4) Flight Characteristics. While all aircraft will have certain minimum requirements for flying qualities, the actual peculiarities and special features of specific airplanes will differ. These particular flight characteristics must be well known and understood by the pilot.

(5) Operating Data. The performance of each specific airplane defines its application to various uses and missions. The handbook operating data must be available at all times to properly plan and execute the flight of an aircraft. Constant reference to the operating data will insure safe and effective operation of the airplane.

Great time and effort are expended in the | preparation of the flight handbook to provide the most exact information, data, and pro­cedures. Diligent study and continuous use of the flight handbook will ensure that the greatest effectiveness is achieved from the airplane while still operating within the inherent capabilities of the design.

[1] GEOPOTENTIAL OF THE TROPOPAUSE Figure 7.7. Standard Altitude Table

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

NAVWEPS 00-80T—80 APPLICATION of aerodynamics TO SPECIFIC PROBLEMS OF FLYING

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

DEAD MAN’S CURVE

A

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.

During takeoff, it is necessary to monitor the performance of the airplane and evaluate the acceleration to insure that the airplane will

achieve the takeoff speed in the specified dis­tance. If it is apparent that the airplane is not accelerating normally or that the airplane or powerplant is not functioning properly, a decision must be made to refuse or continue takeoff. If the decision to refuse takeoff is made early in the takeoff roll, no problem exists because the airplane has not gained much speed and a large portion of runway distance is unused. However, at speeds near the takeoff speed, the airplane has used a large portion of the takeoff distance and the distance required to stop is appreciable. The problem which exists is to define the highest speed attained during takeoff acceleration from which the airplane may be decelerated to a stop on the runway length remaining, i. e., the “refusal speed.’’

The refusal speed will be a function of take­off performance, stopping performance, and the length of available runway. The ideal situation would be to have a runway length which exceeds the total distance required to accelerate to the takeoff speed then decelerate from the takeoff speed. In this case, the refusal speed would exceed the takeoff speed and there would be little concern for the case of refused takeoff. While this may be the case for some instances, the usual case is that the runway length is less than the “accelerate – stop’’ distance and the refusal speed is less than the takeoff speed. A graphical representation of the refused takeoff condition is illustrated in figure 6.12 by a plot of velocity versus dis­tance. At the beginning of the runway, the airplane starts accelerating and the variation of velocity and distance is defined by the takeoff acceleration profile. The deceleration profile describes the variation of velocity with distance where the airplane is brought to a stop at the end of the runway. The inter­section of the acceleration and deceleration profiles then defines the refusal speed and the refusal distance along the runway. Of course, an allowance must be made for the time spent at the refusal speed as the power is reduced and braking action is initiated.

During takeoff, the airplane could be accel­erated to any speed up to the refusal speed, then decelerated to a stop on the runway remaining. Once past the refusal speed, the airplane cannot be brought to a stop on the runway remaining and the airplane is com­mitted to an unsafe stop. If takeoff is refused when above the refusal speed, the only hope is for assistance from the arresting gear, run­way barrier, or an extensive overrun at the end of the runway. This fact points to the need for planning of the takeoff and the require­ment to monitor the takeoff acceleration.

If the refusal speed data are not available, the following equations may be used to ap­proximate the refusal speed and distance:

where

VT = refusal speed Tr = refusal distance

and for the appropriate takeoff configuration,

Vt0= takeoff speed

St0= takeoff distance

VL= landing speed

Ті. = landing distance

Ra = runway length available

These approximate relationships do not ac­count for the time spent at the refusal point and must not be used in lieu of accurate handbook data.

In the case of the single-engine airplane, the pilot must monitor the takeoff performance to recognize malfunctions or lack of adequate ac­celeration prior to reaching the refusal speed. Obviously, it is to advantage to recognize the

possibility of a refused takeoff before exceed­ing the refusal speed. To this end, the pilot must carefully evaluate airplane and power – plant performance and judge the acceleration of the airplane by the use of “line speeds." The accelerated motion of the airplane during takeoff roll will define certain relationships be­tween velocity and distance when the acceler­ation of the airplane is normal. By comparison of predicted and actual speeds at various points along the runway, the pilot can evaluate the acceleration and assess the takeoff perform­ance.

An example of an acceleration profile is shown by the second illustration of figure 6.12, where the variation of velocity and distance is defined for the case of uniformly accelerated motion, i. e., constant acceleration. While the case of uniformly accelerated motion does not correspond exactly to the takeoff performance of all airplanes, it is sufficiently applicable to illustrate the principle of line speeds and ac­celeration checks. If the takeoff acceleration of the airplane were constant, the airplane would develop specific percentages of the take­off speed at specific percentages of the takeoff distance. Representative values from figure 6.12 are as follows:

As an example of this uniformly accelerated motion, the airplane upon reaching the half­way point of takeoff roll would have spent 70.7 percent of the total takeoff time and ac­celerated to 70.7 percent of the takeoff speed. If the airplane has not reached a specific speed at a specific distance, it is obvious that the ac­celeration is below the predicted value and the airplane surely will not achieve the takeoff speed in the specified takeoff distance. There­fore, properly computed line speeds at various points along the runway will allow the pilot to monitor the takeoff performance and recog­nize a deficiency of acceleration. Of course, a deficiency of acceleration must be recognized prior to reaching some point along the runway where takeoff cannot be safely achieved or refused.

The fundamental principles of refusal speeds and line speeds are applicable equally well to single-engine and multiengine airplanes. How­ever, in the case of the multiengine airplane additional consideration must be given to the decision to continue or refuse takeoff when engine failure occurs during the takeoff roll. If failure of one engine occurs prior to reaching the’refusal speed, takeoff should be discon­tinued and the airplane brought to a stop on the remaining runway. If failure of one engine occurs after exceeding the refusal speed, the airplane is committed to continue takeoff with the remaining engines operative or an unsafe refused takeoff. In some cases, the remaining runway may not be sufficient to allow acceler­ation to the takeoff speed and the airplane can neither takeoff or stop on the runway remain­ing. To facilitate consideration of this prob­lem, several specific definitions are necessary.

(1) Takeoff and initial climb speed: A speed,

usually a fixed percentage above the stall speed, at which the airplane will become airborne and best clear obstacles immediately after takeoff. For a particular airplane in the takeoff con­figuration, this speed (in BAS or CAS) is a function of gross weight but in no circumstances should it be less than the minimum directional control speed for the critical asymmetrical power condition. Generally, the takeoff and initial climb speed is referred to as the ‘ V2” speed. •

(2) Critical engine failure speed: A speed

achieved during the takeoff roll at which fail­ure of one engine will require the same distance to continue accelerating with the operative en­gines to accomplish safe takeoff or refuse takeoff and decelerate to a stop utilizing the airplane brakes. At critical engine failure

RUNWAY LENGTH EQUAL TO CRITICAL FIELD LENGTH

speed, the distance necessary to continue take­off with one engine inoperative is equal to the stopping distance. The critical engine failure speed is generally referred to as the "Fj” speed and it is a function of the same factors which determine the takeoff performance, e. g., density altitude, gross weight, temperature, humidity, etc.

(3) Critical field length: The runway length necessary to accelerate with all engines opera­tive to the critical engine failure speed (У0 then continue accelerating to the takeoff and initial climb speed (Fs) with one engine inoperative and achieve safe takeoff or refuse takeoff. By this definition, critical field length describes the minimum length of run­way necessary for safe operation of the multi­engine airplane. Obviously, the critical field length is a function of the same factors affect­ing the takeoff distance of the airplane.

The conditions of Vu and critical field length are illustrated by figure 6.13. The first illustration of figure 6.13 depicts the case where the runway length is equal to the critical field length. In this case, the airplane could accelerate to Fj with all engines opera­tive then either continue takeoff safely with one engine inoperative or refuse takeoff and decelerate to a stop on the remaining runway. For this condition, an engine failure occurring at less than Vt speed dictates that takeoff must be refused because inadequate distance remains to effect a safe takeoff at V3 speed. However, at or below Vi speed, adequate distance re­mains to bring the airplane to a stop. If engine failure occurs at some speed greater than Vt speed, takeoff should be continued because adequate distance remains to accelerate to Fs speed and effect a safe takeoff with one engine inoperative. If engine failure occurs beyond Vi speed, inadequate distance remains to brake the airplane to a stop on the runway.

The second illustration of figure 6.13 depicts the case where the runway length is less than the critical field length. In this case, the term of “Ft” speed is not applicable because of inadequate distance and the refusal speed is less than the minimum speed necessary to continue a safe takeoff with one engine inoper­ative. If engine failure occurs below refusal speed, the takeoff must be refused and adequate distance remains to effect a stop on the runway. If engine failure occurs above refusal speed but below the minimum speed necessary to continue takeoff with one engine inoperative, an accident is inevitable. Within this range of speeds, the airplane cannot effect a safe takeoff at Fa with one engine inoperative or a safe stop on the remaining runway. For this reason, the pilot must properly plan the takeoff and insure that the runway available is equal to or greater than the critical field length. If the runway available is less than the critical field length, there must be sufficient justification for the particular operation be­cause of the hazardous consequences of engine failure between the refusal speed and the minimum speed necessary to continue takeoff with one engine inoperative. Otherwise, the gross weight of the airplane should be reduced in attempt to decrease the critical field length to equal the available runway.

SONIC “BOOMS”

From the standpoint of public relations and the maintaining of friendly public support for Naval Aviation, great care must be taken to prevent sonic booms in populated areas. While the ordinary sonic boom does not carry any potential of physical damage, the disturb­ance must be avoided because of the undesirable annoyance and apprehension. As supersonic flight becomes more commonplace and an ordinary consequence of flying operations, the prevention of sonic booms in populated areas becomes a difficult and perplexing job.

When the airplane is in supersonic flight, the local pressure and velocity changes on the airplane surfaces are coincident with the formation of shock waves. The pressure jump through the shock waves in the immediate vicinity of the airplane surfaces is determined

by the local flow changes at these surfaces. Of course, the strength of the shock waves and the pressure jump through the wave decreases rapidly with distance away from the airplane. While the pressure jump through the shock wave decreases with distance away from the surface, it does not disappear com­pletely and a measurable—but very small— pressure will exist at a considerable distance from the airplane.

Sound is transmitted through the air as a series of very weak pressure waves. In the ordinary range of audible frequencies, the threshold of audibility for intensity of sound is for pressure waves with an approximate R. M.S. value of pressure as low as 0.0000002 psf. Within this same range of frequencies, the threshold of feeling for intensity of sound is for pressure values with an approximate R. M.S. value of pressure of 0.2 to 0,5 psf. Continuous sound at the threshold of feeling is of the intensity to cause painful hearing. Thus, the shock waves generated by an air­plane in supersonic flight are capable of creat­ing audible sound and, in the extreme case, can be of a magnitude to cause considerable dis­turbance. Pressure jumps of 0.02 to 0.3 psf have been recorded during the passage of an airplane in supersonic flight. As a result, the sonic "booms” are the pressure waves generated by the shock waves formed on the airplane in supersonic flight.

The source of sonic booms is illustrated by figure 6.14. When the airplane is in level supersonic flight, a pattern of shock waves is developed which is much dependent on the configuration and flight Mach number of the airplane. At a considerable distance from the airplane, these shock waves tend to combine along two common fronts and extend away from the airplane in a sort of conical surface. The waves decrease in strength with distance away from the airplane but the pressure jump remains of an audible intensity for a consider­able distance from the airplane. If the w-ave extends to the ground or water surface, it will be reflected and attenuated to some extent depending on the character of the reflecting surface. Of course, if this attached wave form is carried across a populated area at the surface, th: population will experience the pressure waves as a sonic boom.

The intensity of the boom will depend on many different factors. The characteristics of the airplane generating the shock waves will be of some importance since a large, high drag, high gross weight airplane in flight at high Mach number will be transferring a greater energy to the air mass. Flight altitude will have an important bearing on boom intensity since at high altitude the pressure jump across a given wave form is much less. In addition, at high altitude a greater distance exists be­tween the generating source of the pressure disturbance and the ground level and the strength of the wave will have a greater dis­tance in which to decay. The ordinary vari­ation of temperature and density plus the natural turbulence of atmosphere will tend to reflect or dissipate the shock wave generated at high altitude. However, in a stable, quies­cent atmosphere, the pressure wave from the airplane in high supersonic flight at high alti­tude may be of an audible magnitude at lateral distances as great as 10 to 30 miles. Thus, supersonic flight over or adjacent to populated areas will produce a sonic boom.

Actually, it is not necessary for an airplane to fly supersonic over or adjacent to a popu­lated area to create a sonic boom. This possibility is shown by the second illustration of figure 6.14 where an airplane decelerates to subsonic from a supersonic dive. As the air­plane slows to subsonic from supersonic speed, the airplane will release the leading bow and tail waves which formed as the airplane accel­erated from subsonic to supersonic speed. The release of these shock waves is analogous to the case where a surface ship slows to below the wave propagation speed and releases the bow wave which then travels out ahead of

the ship. When the airplane slows to sub­sonic, the shock wave travels out ahead of the airplane in a form which is somewhat spheri­cal. Because there are density variations through the shock wave, the shock wave moving ahead of the airplane can cause aber­rations in light waves and it may appear, to the pilot as if a large sheet of clear cellulose or plastic were in front of the airplane. In addition, the density variation and initial shape of the wave leaving the airplane may cause reflection of sunlight which would appear as a sudden, brilliant “flash” to the pilot.

Of course, the wave released by decelerating to subsonic speed can travel out ahead of the airplane and traverse a populated area to cause a sonic boom. The initial direction of the released wave will be the flight path of the airplane at the instant it decelerates to sub­sonic speed. To be sure, the released wave should not be aimed in the direction of a popu­lated area, even if a considerable distance away. There are instances where a released wave has been of an audible magnitude as far as 30 to 40 miles ahead of the point of release. The released pressure wave will be of greatest intensity when created by a large, high drag configuration at low altitude. Since the wave intensity decreases rapidly with distance away from the source, the boom will be of strongest audibility near the point of release.

It should become apparent that sonic booms are a byproduct of supersonic aviation and, with supersonic flight becoming more common­place, the problem is more perplexing. The potential of sonic booms is mostly of the audible nature and nuisance of the disturbance. The damage potential of the ordinary sonic boom is quite small and the principal effects are confined to structures which are extremely brittle, low strength, and have characteristic high residual stresses. In other words, only the extremes of pressure waves generated by airplanes in flight could possibly cause cracked plaster and window glass. Such materials are quite prone to sharp dynamic stresses and, when superimposed on the high residual stresses common to the products and building construction, slight but insignificant damage may result. Actually, the most objectionable feature of the sonic boom is the audibility and the anxiety or apprehension caused by the sharp, loud noise which resembles a blast.

The pressure jump through the shock waves in the immediate vicinity of the airplane is much greater than those common to the audible “booms” at ground level. Thus, airplanes in close formation at supersonic speeds may encounter considerable interference between airplanes. In addition, to eliminate even the most remote possibility of structural damage, a high speed airplane should not make a supersonic pass close to a large air­plane which may have low limit load factor and be prone to be easily disturbed or damaged by a strong pressure wave.

For the majority of airplane configurations and runways conditions, the airplane brakes furnish the most powerful means of deceler­ation. While specific techniques of braking are required for specific situations, there are various fundamentals which are common to all conditions.

Solid friction is the resistance to relative motion of two surfaces in contact. When relative motion exists between the surfaces, the resistance to relative motion is termed “kinetic” or “sliding” friction; when no relative motion exists between the surfaces, the resistance to the impending relative mo­tion is termed “static” friction. The minute discontinuities of the surfaces in contact are able to mate quite closely when relative motion impends rather than exists, so static friction will generally exceed kinetic friction. The magnitude of the friction force between two surfaces will depend in great part on the types of surfaces in contact and the magnitude of force pressing the surfaces together. A convenient method of relating the friction charactersitics of surfaces in contact is a proportion of the friction force to the normal (or perpendicular) force pressing the surfaces together. This proportion defines the coeffi­cient of friction, Ц.

n = F/N

where

n = coefficient of friction (mu)

F = friction force, lbs.

N = normal force, lbs.

The coefficient of friction of tires on a runway surface is a function of many factors. Runway surface condition, rubber composition, tread, inflation pressure, surface friction shearing stress, relative slip speed, etc., all are factors which affect the coefficient of friction. When the tire is rolling along the runway without the use of brakes, the friction force resulting is simple rolling resistance. The coefficient of rolling friction is of an approximate magnitude of 0.015 to 0.030 for dry, hard runway surface.

The application of brakes supplies a torque to the wheel which tends to retard wheel rota­tion. However, the initial application of brakes creates a braking torque but the initial retarding torque is balanced by the increase in friction force which produces a driving or rolling torque. Of course, when the braking torque is equal to the rolling torque, the wheel experiences no acceleration in rotation and the equilibrium of a constant rotational speed is maintained. Thus, the application of brake develops a retarding torque and causes an increase in friction force between the tire and runway surface. A common problem of brak­ing technique is application of excessive brake pressure which creates a braking torque greater than the maximum possible rolling torque. In this case, the wheel loses rotational speed and decelerates until the wheel is stationary and the result is a locked wheel with the tire surface subject to a full slip condition.

The relationship of friction force, normal force, braking torque, and rolling torque is illustrated in figure 6.11.

The effect of slip velocity on the coefficient of friction is illustrated by the graph of figure 6.11. The conditions of zero slip corresponds to the rolling wheel without brake application while the condition of full, 100 percent slip corresponds to the locked wheel where the relative velocity between the tire surface and the runway equals the actual velocity. With the application of brakes, the coefficient of friction increases but incurs a small but meas­urable apparent slip. Continued increase in friction coefficient is obtained until some max­imum is achieved then decreases as the slip increases and approaching the 100 percent slip condition. Actually, the peak value of co­efficient of friction occurs at an incipient skid condition and the relative slip apparent at this point consists primarily of elastic shearing deflection of the tire structure.

When the runway surface is dry, brush – finished concrete, the maximum value for the coefficient of friction for most aircraft tires is on the order of 0.6 to 0.8. Many factors can determine small differences in this peak value of friction coefficient for dry surface conditions. For example, a soft gum rubber composition can develop a very high value of coefficient of friction but only for low values of surface shearing stress. At high values of surface shearing stress, the soft gum rubber will shear or scrub off before high values of friction co­efficient are developed. The higher strength compounds used in the production of aircraft tires produce greater resistance to surface shear and scrubbing but the harder rubber has lower intrinsic friction coefficient. Since the high performance airplane cannot afford the luxury of excessive tire weight or size, the majority of airplane tires will be of relatively hard rubber and will operate at or near the rated load capacities. As a result, there will be little difference between the peak values of friction coefficient for the dry, hard surface runway for the majority of aircraft tires.

If high traction on dry surfaces were the only consideration in the design of tires, the result would be a soft rubber tire of extreme width to create a large footprint and reduce surface shearing stresses, e. g., driving tires on a drag racer. However, such a tire has many other characteristics which are undesirable such as high rolling friction, large size, poor side force characteristics, etc.

When the runway has water or ice on the surface, the maximum value for the coefficient of friction is reduced greatly below the value obtained for the dry runway condition. When water is on the surface, the tread design be­comes of greater importance to maintain con­tact between the rubber and the runway and prevent a film of water from lubricating the surfaces. When the rainfall is light, the peak value for friction coefficient is on the order of 0.5, With heavy rainfall it is more likely that sufficient water will stand to form a liquid film between the tire and the runway. In this case, the peak coefficient of friction rarely exceeds 0.3. In some extreme conditions, the tire may simply plane along the water without contact of the runway and the coefficient of friction is much lower than 0.3. Smooth, clear ice on the runway will cause extremely low values for the coefficient of friction. In such a condition, the peak value for the co­efficient of friction may be on the order of 0.2 or 0.15-

Note that immediately past the incipient skidding condition the coefficient of friction decreases with increased slip speed, especially for the wet or icy runway conditions. Thus, once skid begins, a reduction in friction force and rolling torque must be met with a reduc­tion in braking torque, otherwise the wheel will decelerate and lock. This is an important factor to consider in braking technique because the skidding tire surface on the locked wheel produces considerably less retarding force than when at the incipient skid condition which causes the peak coefficient of friction. If the wheel locks from excessive braking, the sliding tire surface produces less than the maximum retarding force and the tires become relatively incapable of developing any significant side force. Stop distance will increase and it may be difficult—if not impossible—to control the airplane when full slip is developed. In addi­tion, at high rolling velocities on the dry sur­face runway, the immediate problem of a skid­ding tire is not necessarily the loss of retard­ing force but the imminence of tire failure. The pilot must insure that the application of brakes does not produce some excessive braking torque which is greater than the maximum rolling torque and particular care must be taken when the runway conditions produce low values of friction coefficient and when the normal force on the braking surfaces is small. When it is difficult to perceive or distinguish a skidding condition, the value of an antiskid or auto­matic braking system will be appreciated.

BRAKING TECHNIQUE. It must be clearly distinguished that the techniques for minimum stop distance may differ greatly from the techniques required to minimize wear and tear on the tires and brakes. For the majority of airplane configurations, brakes will provide the most important source of deceleration for all but the most severe of icy runway condi­tions. Of course, aerodynamic drag is very durable and should be utilized to decelerate the airplane if the runway is long enough and the drag high enough. Aerodynamic drag will be of importance only for the initial 20 to 30 per­cent of speed reduction from the point of touch­down, At speeds less than 60 to 70 percent of the landing speed, aerodynamic drag is of little consequence and brakes will be the principal source of deceleration regardless of the runway surface. For the conditions of minimum land­ing distance, aerodynamic drag will be a prin­cipal source of deceleration only for the initial portion of landing roll for very high drag con­figurations on very poor runway conditions. These cases are quite limited so considerable importance must be assigned to proper use of the brakes to produce maximum effectiveness.

In order to provide the maximum possible retarding force, effort must be directed to pro­duce the maximum normal force on the braking surfaces. (See figure 6.11.) The pilot will be able to influence the normal force on the brak­ing surfaces during the initial part of the land­ing roll when dynamic pressure is large and aerodynamic forces and moments are of conse­quence. During this portion of the landing roll the pilot can control the airplane lift and the distribution of normal force to the landing gears.

First to consider is that any positive lift will support a part of the airplane weight and reduce the normal force on the landing gear. Of course, for the purposes of braking friction, it would be to advantage to create negative lift but this is not the usual capability of the air­plane with the tricycle landing gear. Since the airplane lift may be considerable immedi­ately after landing, retraction of flaps or ex­tension of spoilers immediately after touch­down will reduce the wing lift and increase the normal force on the landing gear. With the retraction of flaps, the reduced drag is more than compensated for by the increased braking friction force afforded by the increased normal force on the braking surfaces.

A second possible factor to control braking effectiveness is the distribution of normal force to the landing gear surfaces. The nose wheel of the tricycle landing gear configuration usu­ally has no brakes and any normal force dis­tributed to this wheel is useful only for pro­ducing side force for control of the airplane. Under conditions of deceleration, the nose – down pitching moment created by the friction force and the inertia force tends to transfer a significant amount of normal force to nose wheel where it is unavailable to assist in creating friction force. For the instant after landing touchdown, the pilot may control this condition to some extent and regain or increase the normal force on the main wheels. After touchdown, the nose is lowered until the nose wheel contacts the runway then brakes are applied while the stick is eased back with­out lifting the nose wheel back off the runway. The effect is to minimize the normal force on the nose wheel and increase the normal force on braking surfaces. While the principal effect is to transfer normal force to the main wheels, there may be a significant increase in normal force due to a reduction in net lift, i. e., tail download is noticeable. This reduction in net lift tends to be particular to tailless or short coupled airplane configurations.

The combined effect of flap retraction and aft stick is a significant increase in braking friction force. Of course, the flaps should not be retracted while still airborne and aft stick should be used just enough without lifting the nosewheel off the runway. These tech­niques are to no avail if proper use of the

brakes does not produce the maximum coeffi­cient of friction. The incipient skid condi­tion will produce the maximum coefficient of friction but this peak is difficult to recognize and maintain without an antiskid system. Judicious use of the brakes is necessary to obtain the peak coefficient of friction but not develop a skid or locked wheel which could cause tire failure, loss of control, or consider­able reduction in the friction coefficient.

The capacity of the brakes must be sufficient to create adequate braking torque and produce the high coefficient of friction. In addition, the brakes must be capable of withstanding the heat generated without fading or losing effectiveness. The most critical requirements of the brakes occur during landing at the maximum allowable landing weight.

TYPICAL ERRORS OF BRAKING TECH­NIQUE. Errors in braking technique are usu­ally coincident with errors of other sorts. For example, if the pilot lands an airplane with excessive airspeed, poor braking technique could accompany the original error to produce an unsafe situation. One common error of of braking technique is the application of braking torque in excess of the maximum possible rolling torque. The result will be that the wheel decelerates and locks and the skid reduces the coefficient of friction, lowers the capability for side force, and enhances the possibility of tire failure. If maximum brak­ing is necessary, caution must be used to modulate the braking torque to prevent lock­ing the wheel and causing a skid. On the other hand, maximum coefficient of friction is obtained at the incipient skidding condition so sufficient brake torque must be applied to produce maximum friction force. Intermittent braking serves no useful purpose when the objective is maximum deceleration because the periods between brake application produce only slight or negligible cooling. Brake should be applied smoothly and braking torque modulated at or near the peak value to insure that skid does not develop.

One of the important factors affecting the landing roll distance is landing touchdown speed. Any excess velocity at landing causes a large increase in the minimum stop distance and it is necessary that the pilot control the landing precisely so to land at the appropriate speed. When landing on the dry, hard surface runway of adequate length, a tendency is to take advantage of any excess runway and allow the airplane to touchdown with excess speed. Of course, such errors in technique cannot be tolerated and the pilot must strive for precision in all landings. Immediately after touchdown, the airplane lift may be considerable and the normal force on the braking surfaces quite low. Thus, if excessive braking torque is applied, the wheel may lock easily at high speeds and tire failure may take place suddenly.

Landing on a wet or icy runway requires judicious use of the brakes because of the re­duction in the maximum coefficient of friction. Because of reduction in the maximum attain­able value of the coefficient of friction, the pilot must anticipate an increase in the mini­mum landing distance above that applicable for the dry runway conditions. When there is considerable water or ice on the runway, an increase in landing distance on the order of 40 to 100 percent must be expected for similar conditions of gross weight, density altitude, wind, etc. Unfortunately, the conditions likely to produce poor braking action also will cause high idle thrust of the turbojet engine and the extreme case (smooth, glazed ice or heavy rain) may dictate shutting down the engine to effect a reasonable stopping distance.

During formation flying and inflight refuel­ing, airplanes in proximity to one another will produce a mutual interference of the flow pat­terns and alter the aerodynamic characteristics of each airplane. The principal effects of this interference must be appreciated since certain factors due to the mutual interference may enhance the possibility of a collision.

One example of interference between air­planes in flight is shown first in figure 6.10 with the effect of lateral separation of two airplanes flying in line abreast. A plane of symmetry would exist halfway between two identical air­planes and would furnish a boundary of flow across which there would be no lateral com­ponents of flow. As the two airplane wing tips are in proximity, the effect is to reduce the strength of the tip or trailing vortices and re­duce the induced velocities in the vicinity of wing tip. Thus, each airplane will experience a local increase in the lift distribution as the tip vortices are reduced and a rolling moment is developed which tends to roll each airplane away from the other. This disturbance may provide the possibility of collision if other air­planes are in the vicinity and there is delay in control correction or overcontrol. If the wing tips are displaced in a fore-and-aft direction, the same effect exists but generally it is of a lower magnitude.

The magnitude of the interference effect due to lateral separation of the wing tips depends on the proximity of the wi. ig tips and the ex­tent of induced flow. This implies that the interference vr ’ ‘ e greatest when the tips are very close aixa the airplanes are operating at high lift coefficients. An interesting ramifi­cation of this effect is that several airplanes in line abreast with the wing tips quite close will experience a reduction in induced drag.

An indirect form of interference can be en­countered from the vortex system created by a preceding airplane along the intended flight path. The vortex sheet rolls up a considerable distance behind an airplane and creates consid­erable turbulence for any closely following air­plane. This wake can prove troublesome if air­planes taking off and landing are not provided adequate separation. The rolled-up vortex sheet will be strongest when the preceding air­planes is large, high gross weight, and operat­ing at high lift coefficients. At times this tur­bulence may be falsely attributed to propwash or jet wash.

Another important form of direct inter­ference is common when the two airplanes are in a trail position and stepped down. As shown in figure 6.10, the single airplane in flight de­velops upwash ahead of the wing and down – wash behind and any restriction accorded the flow can alter the distribution and magnitude of the upwash and downwash. When the trailing airplane is in dose proximity aft and below the leading airplane a mutual interference takes place between the two airplanes. The leading airplane above will experience an effect which would be somewhat similar to encountering ground effect, i. e., a reduction in induced drag, a reduction in downwash at the tail, and a change in pitching moment nose down. The trailing airplane below will experience an effect which is generally the opposite of the airplane above. In other words* the airplane below will experience an increase in induced drag, an increase in downwash at the tail, and a change in pitching moment nose up. Thus, when the airplanes are in close proximity, a definite collision possibility exists because of the trim change experienced by each airplane. The magnitude of the trim change is greatest when the airplanes are operating at high lift coefficients, e. g., low speed flight, and when the airplanes are in close proximity.

In formation flying, this sort of interference must be appreciated and anticipated. In cross­ing under another airplane, care must be taken to anticipate the trim change and adequate clearance must be maintained, other­wise a collision may result. The pilot of the leading aircraft will know of the presence of the trailing airplane by the trim change experienced. Obviously, some anticipation is necessary and adequate separation is necessary to prevent a disturbing magnitude of the trim change. In a close diamond formation the leader will be able to ‘‘feel” the presence of the slot man even though the airplane is not within view. Obviously, the slot man will have a difficult job during formation maneuvers because of the unstable trim changes and greater power changes required to hold position.

A common collision problem is the case of an airplane with a malfunctioning landing gear. If another airplane is called to inspect the malfunctioning landing gear, great care must be taken to maintain adequate separation and preserve orientation. Many instances such as this have resulted in a collision when the pilot of the trailing airplane became dis­oriented and did not maintain adequate sepa­ration.

During inflight refueling, essentially the same problems of interference exist. As the receiver approaches the tanker from behind and below, the receiver will encounter the downwash from the tanker and require a slight, gradual increase in power and pitch attitude to continue approach to the receiving position. While the receiver may not be visible to the pilot of the tanker, he will anticipate the receiver coming into position by the slight reduction in power required and nose down change in pitching moment. Ade­quate clearance and1 proper position must be maintained by the pilot of the receiver for a collision possibility is enhanced by the rela­tive positions of the airplanes. A hazardous condition exists if the pilot of the receiver has excessive speed and runs under the tanker in close proximity. ‘ The trim change expe­rienced by both airplanes may be large and unexpected and it may be difficult to avoid a collision. ‘

In addition to the forms of interference previously mentioned, there exists the possi­bility of strong interference between airplanes in supersonic flight. In this case, the shock waves from one airplane may strongly affect the pressure distribution and rolling, yawing, and pitching moments of an adjacent air­plane. It is difficult to express general rela­tionships of the effect except that magnitude of the effects will be greatest when in close proximity at low altitude and high q. General­ly, the trailing airplane will be most affected.

When an airplane in flight nears the ground (or water) surface, a change occurs in the three dimensional flow pattern because the local airflow cannot have a vertical component at the ground plane. Thus, the ground plane will furnish a restriction to the flow and alter the wing upwash, downwash, and tip vortices. These general effects due to the presence of the ground plane are referred to as “ground effect.”

AERODYNAMIC INFLUENCE OF GROUND EFFECT. While the aerodynamic characteristics of the tail and fuselage are altered by ground effects, the principal effects due to proximity of the ground plane are the changes in the aerodynamic characteristics of the wing. As the wing encounters ground effect and is maintained at a constant lift coefficient, there is a reduction in the upwash, downwash, and the tip vortices. These effects are illustrated by the sketches of figure 6.9. As a result of the reduced tip vortices, the wing in the presence of ground effect will behave as if it were of a greater aspect ratio. In other words, the induced velocities due to the tip (or trailing) vortices will be reduced and the wing will incur smaller values of induced drag coefficient, Cu., and induced angle of attack, on, for any specific lift coefficient, CL.

In order for ground effect to be of a signifi­cant magnitude, the wing must be quite close to the ground plane. Figure 6.9 illustrates one of the direct results of ground effect by the variation of induced drag coefficient with wing height above the ground plane for a representative unswept wing at constant lift coefficient. Notice that the wing must be quite close to the ground for a noticeable reduction in induced dfag. When the wing is at a height equal to the span (hjb=1.0), the reduction in induced drag is only 1.4 percent. However, when the wing is at a height equal to one-fourth the span (hjb = 0.25), the reduction in induced drag is 23.5

percent and, when the wing is at a height equal to one-tenth the span (hjb=0.1), the reduction in induced drag is 47.6 percent. Thus, a large reduction in induced drag will take place only when the wing is very close to the ground. Because of this variation, ground effect is most usually recognized during the liftoff of takeoff or prior to touchdown on landing.

The reduction of the tip or trailing vortices due to ground effect alters the spanwise lift distribution and reduces the induced angle of attack. In this case, the wing will require a lower angle of attack in ground effect to produce the same lift coefficient. This effect is illustrated by the lift curves of figure 6.9 which show that the airplane in ground effect will develop a greater slope of the lift curve. For the wing in ground effect, a lower angle of attack is necessary to produce the same lift coefficient or, if a constant angle of attack is maintained, an increase in lift coefficient will result.

Figure 6.9 illustrates the manner in which ground effect will alter the curve of thrust re­quired versus velocity. Since induced drag predominates at low speeds, the reduction of induced drag due to ground effect will cause the most significant reduction of thrust re­quired (parasite plus induced drag) only at low speeds. At high speeds where parasite drag predominates, the induced drag is but a small part of the total drag and ground effect causes no significant change in thrust re­quired. Because ground effect involves the induced effects of airplane when in close prox­imity to the ground, its effects are of greatest concern during the takeoff and landing. Ordi­narily, these are the only phases of flight in which the airplane would be in close proximity to the ground.

GROUND EFFECT ON SPECIFIC FLIGHT CONDITIONS. The overall influence of ground effect is best realized by assuming that the airplane descends into ground effect while maintaining a constant lift coefficient and, thus, a constant dynamic pressure and equiva­lent airspeed. As the airplane descends into ground effect, the following, effects will take place:

(1) Because of the reduced induced angle of attack and change in lift distribution, a smaller wing angle of attack will be required to produce the same lift coefficient. If a constant pitch attitude is maintained as ground effect is encountered, an increase in lift coefficient will be incurred.

(2) The reduction in induced flow due to ground effect causes a significant reduction in induced drag but causes no direct effect on parasite drag. As a result of the reduction in induced drag, the thrust required at low speeds will be reduced.

(3) The reduction in downwash due to ground effect will produce a change in longi­tudinal stability and trim. Generally, the reduction in downwash at the horizontal tail increases the contribution to static longi­tudinal stability. In addition, the reduction of downwash at the tail usually requires a greater up elevator to trim the airplane at a specific lift coefficient. For the conven­tional airplane configuration, encountering ground effect will produce a nose-down change in pitching moment. Of course, the increase in stability and trim change associ­ated with ground effect provide a critical re­quirement of adequate longitudinal control power for landing and takeoff.

(4) Due to the change in upwash, down – wash, and tip vortices, there will be a change in position error of the airspeed system, as­sociated with ground effect. In the majority of cases, ground effect will cause an increase in the local pressure at the static source and produce a lower indication of airspeed and altitude.

During the landing phase of flight, the effect of proximity to the ground plane must be understood and appreciated. If the airplane is brought into ground effect with a constant angle of attack, the airplane will experience

an increase in lift coefficient and reduction in thrust required. Hence, a “floating” sensa­tion may be experienced. Because of the re­duced drag and power-off deceleration in ground effect, any excess speed at the point of flare may incur a considerable “float” distance. As the airplane nears the point of touchdown on the approach, ground effect will be most realized at altitudes less than the wing span. An exact appreciation of the ground effect may be obtained during a field approach with the mirror landing system furnishing an exact reference of the flight path. During the final phases of the field approach as the airplane nears the ground plane, a reduced power setting is necessary or the reduced thrust re­quired would allow the airplane to climb above the desired glide path. During ship­board operations, ground effect will be delayed until the airplane passes the edge of the deck and the reduction in power setting that is common to field operations should not be encountered. Thus, a habit pattern should not be formed during field landings which would prove dangerous during carrier oper­ations.

An additional factor to consider is the aero­dynamic drag of the airplane during the land­ing roll. Because of the reduced induced drag when in ground effect, aerodynamic braking will be of greatest significance only when partial stalling of the wing can be accom­plished. The reduced drag when in ground effect accounts for the fact that the brakes are the most effective source of deceleration for the majority of airplane configurations.

During the takeoff phase of flight ground effect produces some important relationships. Of course, the airplane leaving ground effect encounters just the reverse of the airplane entering ground effect, i. e., the airplane leaving ground effect will (1) require an increase in angle of attack to maintain the same lift coefficient, (2) experience an increase in in­duced drag and thrust required, (3) experience a decrease in stability and a nose-цр change in moment, and (4) usually a reduction in static source pressure and increase in indicated air­speed. These general effects should point out the possible danger in attempting takeoff prior to achieving the recommended takeoff speed. Due to the reduced drag in ground effect the airplane may seem capable of takeoff below the recommended speed. However, as the airplane rises out of ground effect with a deficiency of speed, the greater induced drag may produce marginal initial climb perform­ance. In the extreme conditions such as high gross weight, high density altitude, and high temperature, a deficiency of airspeed at takeoff may permit the airplane to become airborne but be incapable of flying out of ground effect. In this case, the airplane may become airborne initially with a deficiency of speed, but later settle back to the runway. It is imperative that no attempt be made to force the airplane to become airborne with a deficiency of speed; the recommended takeoff speed is necessary to provide adequate initial climb performance. In fact, ground effect can be used to advantage if no obstacles exist by using the reduced drag to improve initial acceleration.

The results of the airplane leaving ground effect can be most easily realized during the deck launch of a heavily loaded airplane. As the airplane moves forward and passes over the edge of the deck, whatever ground effect exists will be lost immediately. Thus, proper rota­tion of the airplane will be necessary to main­tain the same lift coefficient and the increase in induced drag must be expected.

The rotor of the helicopter experiences a similar restraint of induced flow when in prox­imity to the ground plane. Since the induced rotor power required will predominate at low flight speeds, ground effect will produce a con­siderable effect on the power required at low speeds. During hovering and flight at low speeds, the elevation of the rotor above the ground plane will be an important factor de­termining the power required for flight.

The range of the reciprocating powered air­plane can be augmented by the use of ground effect. When the airplane is close to the ground or water surface the reduction of in­duced drag increases the maximum lift-drag ratio and causes a corresponding increase in range. Of course, the airplane must be quite close to the surface to obtain a noticeable in­crease in (LfU)max and range. The difficulty in holding the airplane at the precise altitude without contacting the ground or water will preclude the use of ground effect during ordi­nary flying operations. The use of ground effect to extend range should be reserved as a final measure in case of emergency. Because of the very detrimental effect of low altitude on the range of the turbojet, ground effect will not be of a particular advantage in an attempt to augment range.

The most outstanding examples of the use of ground effect are shown in the cases of multi­engine airplanes with some engines inoperative. When the power loss is quite severe, the air­plane may not be capable of sustaining altitude and will descend. As ground effect is en­countered, the reduced power required may allow the airplane to sustain flight at extremely low altitude with the remaining powerplants functioning. In ground effect, the recipro­cating powered airplane will encounter a greater (_L/D’)max which occurs at a lower air­speed and power required and the increase in range may be quite important during emer­gency conditions.

In the case of the single-engine airplane, powerplant failure leaves only the alternatives of effecting a successful power-off landing or abandoning the airplane. In the case of the multiengine airplane, the failure of a power- plant does not necessarily constitute a disaster since flight may be continued with the remain­ing powerplants functioning. However, the performance of the multiengine airplane with a powerplant inoperative may be critical for certain conditions of flight and specific tech­niques and procedures must be observed to obtain adequate performance.

The effect of a powerplant failure on the multiengine turbojet airplane is illustrated by the first chart of figure 6.8 with the variation of required and available thrust with velocity. If half of the airplane powerplants are inoper­ative, e. g., single-engine operation of a twin – engine airplane, the maximum thrust available at each velocity is reduced to half that avail­able prior to the engine failure. The variation of thrust required with velocity may be affected by the failure of a powerplant in that there may be significant increases in drag if specific procedures are not followed. The inoperative powerplant may contribute addi­tional drag and the pilot must insure that the additional drag is held to a minimum. In the case of the propeller powered airplane, the propeller must be feathered, cowl flaps closed, etc., as the increased drag will detract con­siderably from the performance.

The principal effects of the reduced available thrust are pointed out by the illustration of figure 6.8. Of course, the lower available thrust will reduce the maximum level flight speed but of greater importance is the reduc­tion in excess thrust. Since the acceleration and climb performance is a function of the excess thrust and power, the failure of a power – plant will be most immediately appreciated in this area of performance. As illustrated in figure 6.8, loss of one-half the maximum avail­able thrust will reduce the excess thrust to less than half the original value. Since some thrust is required to sustain flight, the excess which remains to accelerate and climb the airplane may be greatly reduced. The most critical conditions will exist when various factors combine to produce a minimum of excess thrust or power when engine failure occurs. Thus, critical conditions will be com­mon to high gross weight and high density altitude (and high temperatures in the case of the turbine powered airplane) as each of these factors will reduce the excess thrust at any specific flight condition.

The asymmetrical power condition which results when a powerplant fails can provide critical control requirements. First consid­eration is due the yawing moment produced by the asymmetrical power condition. Ade­quate directional control will be available only when the airplane speed is greater than the minimum directional control speed. Thus, the pilot must insure that the flight speed never falls below the minimum directional control speed because the application of maximum power on the functioning powerplants will produce an uncontrollable yaw if adequate directional control is unavailable. A second consideration which is due the propeller powered airplane involves the rolling moments caused by the slipstream velocity. Asym­metrical power on the propeller airplane will create a dissymmetry of the slipstream veloc­ities on the wing and create rolling moments which must be controlled. These slipstream induced rolling moments will be greatest at high power and low velocity and the pilot must be sure of adequate lateral control, especially for the crosswind landing.

The effect of an engine failure on the remain­ing range and endurance is specific to the air­plane type and configuration. If an engine fails during optimum cruise of the turbojet airplane, the airplane must descend and experi­ence a loss of range. Since the turbojet air­plane is generally overpowered at QLjD)max, a loss of a powerplant will not cause a signi­ficant change in maximum endurance. If an engine fails during cruise of a reciprocating powered airplane, there will be a significant loss of range only if the maximum range condi­tion cannot be sustained with the remaining powerplants operating within the cruise power rating. If a power greater than the maximum cruise rating is necessary to sustain cruise, the specific fuel consumption increases and causes a reduction of range. Essentially the same relationship exists regarding maximum endur­ance of the reciprocating powered airplane.

When critical conditions exist due to failure of a powerplant, the pilot must appreciate the reduced excess thrust and operate the airplane within specific limitations. If the engine-out performance of the airplane is marginal, the pilot must be aware of the very detrimental effect of steep turns. Due to the increased load factor in a coordinated turn, there will be an increase in stall speed and—of greater import­ance to engine-out performance—an increase in induced drag. The following table illus­trates the effect of bank angle on stall speed and induced drag.

The previous table of values illustrates the fact that coordinated turns with less than 15° of bank cause no appreciable effect on stall speed or induced drag. However, note that 30° of bank will increase the induced drag by 33.3 percent. Under critical conditions, such an in­crease in induced drag (and, hence, total drag) would be prohibitive causing the airplane to descend rather than climb. The second graph of figure 6.7 illustrates the case where the steep turn causes such a large increase in required thrust that a deficiency of thrust exists. When­ever engine failure produces critical perform­ance conditions it is wise to limit all turns to 15° of bank wherever possible.

Another factor to consider in turning flight is the effect of sideslip. If the turn is not coor­dinated to hold sideslip to a minimum, addi­tional drag will be incurred due to the sideslip.

The use of the flaps and landing gear can greatly affect the performance of the multi­engine airplane when a powerplant is inopera­tive. Since the extension of the landing gear and flaps increases the parasite drag, maximum performance of the airplane will be obtained with airplane in the clean configuration. In certain critical conditions, the extension of the landing gear and full flaps may create a defi­ciency of thrust at any speed and commit the airplane to descend. This condition is illus­trated by the second graph of figure 6.8. Thus, judicious use of the flaps and landing gear is necessary in the case of an engine failure.

In the case of engine failure immediately after takeoff, it is important to maintain air­speed in excess of the minimum directional con­trol speed and accelerate to the best climb speed. After the engine failure, it will be fa­vorable to climb only as necessary to clear obstacles until the airplane reaches the best climb speed. Of course, the landing gear should be retracted as soon as the airplane is airborne to reduce para­site drag and, in the case of the propeller pow­ered airplane, it is imperative that the wind milling propeller be feathered. The flaps should be retracted only as rapidly as the increase in

THRUST REQ’D LANDING CONFIGURATION WING LEVEL FLIGHT

THRUST REQ’D CLEAN

configuration

BUT TURNING FLIGHT

REDUCED THRUST AVAILABLE DUE TO ENGINE FAILURE

VELOCITY, KNOTS

airspeed will allow. If full flap deflection is utilized for takeoff it is important to recall that the last 50 percent of flap deflection creates more than half the total drag increase but less than half the total change in CLmax – Thus, for some configurations of airplanes, a greater re­duction in drag may be accomplished by partial retraction of the flaps rather than retraction of the landing gear. Also, it is important that no steep turns be attempted because of the unde­sirable increase in induced drag.

During the landing with an engine inopera­tive, the same fundamental precautions must be observed as during takeoff, i. e., minimum directional control speed must be maintained (or exceeded), no steep turns should be at­tempted, and the extension of the flaps and landing gear must be well planned. In the case of a critical power condition it may be neces­sary to delay the extension of the landing gear and full flaps until a successful landing is as­sured. If a waveoff is necessary, maximum per­formance will be obtained cleaning up the air­plane and accelerating to the best climb speed before attempting any gain in altitude.

At all times during flight with an engine inoperative, the pilot must utilize the proper techniques for control of airspeed and altitude, e. g., for the conditions of steady flight, angle of attack is the primary control of airspeed and excess power is the primary control of rate of climb. For example, if during approach to landing the extension of full flaps and landing gear creates a deficiency of power at all speeds, the airplane will be committed to descend. If the approach is not properly planned and the airplane sinks below the desired glide path, an increase in angle of attack will only allow the airplane to fly more slowly and descend more rapidly. An attempt to hold altitude by increased angle of attack when a power deficiency exists only causes a continued loss of airspeed. Proper procedures and technique are an absolute necessity for safe flight when an engine failure occurs.

Without exception, the formation of ice or frost on the surfaces of an airplane will cause a detrimental effect on aerodynamic performance. The ice or frost formation on the airplane sur­faces will alter the aerodynamic contours and affect the nature of the boundary layer. Of course, the most important surface of the air­plane is the wing and the formation of ice or frost can create significant changes in the aero­dynamic characteristics.

LEADING EDGE ICE FORMATION

A large formation of ice on the leading edge of the wing can produce large changes in the local contours and severe local pressure gra­dients. The extreme surface roughness common to some forms of ice will cause high surface friction and a considerable reduction of bound­ary layer energy. As a result of these effects, the ice formation can produce considerable in­crease in drag and a large reduction in maxi­mum lift coefficient. Thus, the ice formation will cause an increase in power required and stall speed. In addition, the added weight of the ice formation on the airplane will provide an undesirable effect. Because of the detri­mental effects of ice formation, recommended anti-icing procedures must be followed to preserve the airplane performance.

The effect of frost is perhaps more subtle than the effect of ice formation on the aero­dynamic characteristics of the wing. The ac­cumulation of a hard coat of frost on the wing upper surface will provide a surface texture of considerable roughness. While the basic shape and aerodynamic contour is unchanged, the increase in surface roughness increases skin – friction and reduces the kinetic energy of the boundary layer. As a result, there will be an increase in drag but, of course, the magnitude of drag increase will not compare with the considerable increase due to a severe ice forma­tion. The reduction of boundary layer kinetic energy will cause incipient stalling of the wing,

i. e., separation will occur at angles of attack and lift coefficients lower than for the clean, smooth wing. While the reduction in due to frost formation ordinarily is not as great as that due to ice formation, it is usually un­expected because it may be thought that large changes in the aerodynamic shape (such as due to ice) are necessary to reduce CL. How­ever, the kinetic energy of the boundary layer is an important factor influencing separation of the airflow and this energy is reduced by an increase in surface roughness.

The general effects of ice and frost formation on the lift characteristics is typified by the il* lustration of figure 6.7.

The effect of ice or frost on takeoff and land­ing performance is of great importance. The effects are so detrimental to the landing and takeoff that no effort should be spared to keep the airplane as free as possible from any ac­cumulation of ice or frost. If any ice remains on the airplane as the landing phase approaches it must be appreciated that the ice formation will have reduced Ct and incurred an increase in stall speed. Thus, the landing speed will be greater. When this effect is coupled with the possibility of poor braking action during the landing roll, a critical situation can exist. It is obvious that great effort must be made to prevent the accumulation of ice during flight.

In no circumstances should a formation of ice or frost be allowed to remain on the airplane wing surfaces prior to takeoff. The undesir­able effects of ice are obvious but, as previously mentioned, the effects of frost are more subtle. If a heavy coat of hard frost exists on the wing upper surface, a typical reduction in CL would cause a 5 to 10 percent increase in the airplane stall speed. Because of this magnitude of effect, the effect of frost on takeoff per­formance may not be realized until too late. The takeoff speed of an airplane is generally some speed 5 to 25 percent greater than the stall speed, hence the takeoff lift coefficient will be value from 90 to 65 percent of. Thus, it is possible that the airplane with frost cannot become airborne at the specified take­off speed because of premature stalling. Even if the airplane with frost were to become air­borne at the specified takeoff speed, the air­plane could have insufficient margin of air­speed above stall and turbulence, gusts, turning flight could produce incipient or con plete stalling of the airplane.

The increase in drag during takeoff roll due to frost or ice is not considerable and there will not be any significant effect on the initial acceleration during takeoff. Thus, the effect of frost or ice will be most apparent during the later portions of takeoff if the airplane is un­able to become airborne or if insufficient margin above stall speed prevents successful initial climb. In no circumstances should a formation of ice or frost be allowed to remain on the air­plane wing surfaces prior to takeoff.

The gliding performance of an airplane is of special interest for the single-engine airplane
in the case of powerplant failure or malfunc­tion. When a powerplant failure or malfunc­tion occurs, it is usually of interest to obtain a gliding flight path which results in the mini­mum glide angle. The minimum glide angle will produce the greatest proportion of glide distance to altitude loss and will result in maximum glide range or minimum expendi­ture of altitude for a specific glide distance.

GLIDE ANGLE AND LIFT-DRAG RATIO. In the study of climb performance, the forces acting on the airplane in a steady climb (or glide) produce the following relationship:

sin y~

where

7 ^ angle of climb, degrees T= thrust, lbs.

D=drag, lbs.

W^lbs.

In the case of power-off glide performance, the thrust, T, is zero and the relationship reduces to:

sin y= —

By this relationship it is evident that the mini­mum angle of glide—or minimum negative climb angle—is obtained at the aerodynamic conditions which incur the minimum total drag. Since the airplane lift is essentially equal to the weight, the minimum angle of glide will be obtained when the airplane is operated at maximum lift-drag ratio, (L/D)mo;e. When the angle of glide is relatively small, the ratio of glide distance to glide altitude is numeri­cally equal to the airplane lift-drag ratio.

glide distance, ft. glide altitude, ft.

glide ratio = (L/D)

Figure 6.6 illustrates the forces acting on the airplane in a power-off glide. The equilibrium of the steady glide is obtained when the sum­mation of forces in the vertical and horizontal directions is equal to zero.

In order to obtain maximum glide ratio, the airplane must be operated at the angle of at­tack and lift coefficient which provide maxi­mum lift-drag ratio. The illustration of figure 6.6 depicts a variation of lift-drag ratio, LjD, with lift coefficient, CL, for a typical airplane in the clean and landing configurations. Note that {LjD)max for each configuration will occur at a specific value of lift coefficient and, hence, a specific angle of attack. Thus, the maximum glide performance of a given airplane configu­ration will be unaffected by gross weight and altitude when the airplane is operated at (L(D’)max. Of course, an exception occurs at very high altitudes where compressibility ef­fects may alter the aerodynamic characteristics. The highest value of (LjD) will occur with the airplane in the clean configuration. As the airplane is changed to the landing configura­tion, the added parasite drag reduces (LjD)max and the CL which produces (L/D)imi will be in­creased. Thus, the best glide speed for the landing configuration generally will be less than the best glide speed for the clean configuration.

The power-off glide performance may be appreciated also by the graph of rate of descent versus velocity shown in figure 6.6. When a straight line is drawn from the origin tangent to the curve, a point is located which produces the maximum proportion of velocity to rate of descent. Obviously, this condition provides maximum glide ratio. Since the rate of descent is proportional to the power required, the points of tangency define the aerodynamic condition of (L/D),mx.

FACTORS AFFECTING GLIDE PER­FORMANCE, In order to obtain the mini­mum glide angle through the air, the airplane must be operated at {LjD)maz. The subsonic (LjD)max of a given airplane configuration will occur at a specific value of lift coefficient and angle of attack. However, as can be noted from the curves of figure 6.6, small deviations from the optimum CL will not cause a drastic reduction of {LjD’) and glide ratio. In fact, a 5 percent deviation in speed from the best glide speed will not cause any significant reduction of glide ratio. This is fortunate and allows the specifying of convenient glide speeds which will be appropriate for a range of gross weights at which power-off gliding may be encoun­tered, e. g., small quantities of fuel remaining.

An attempt to stretch a glide by flying at speeds above or below the best glide speed will prove futile. As shown by the illustration of figure 6.6, any CL above or below the optimum will produce a lift-drag ratio less than the maximum. If the airplane angle of attack is increased above the value for {LJD)mazt a tran­sient reduction in rate of descent will take place but this process must be reserved for the land­ing phase. Eventually, the steady-state condi­tions would be achieved and the increased angle of attack would incur a lower airspeed and a reduction in {LjD) and glide ratio.

The effect of gross weight on glide performance may be difficult to appreciate. Since (LjD)maz of a given airplane configuration will occur at a specific value of CL, the gross weight of the air­plane will not affect the glide ratio if the air­plane is operated at the optimum CL. Thus, two airplanes of identical aerodynamic con­figuration but different gross weight could glide the same distance from the same altitude. Of course, this fact would be true only if both airplanes are flown at the specific CL to produce (LjD)№>iU.. The principal difference would be that the heavier airplane must fly at a higher airspeed to support the greater weight at the optimum CL. In addition, the heavier airplane flying at the greater speed along the same flight path would develop a greater rate of descent.

The relationship which exists between gross weight and velocity for a particular Cb is as follows:

CconstantC‘)

where

Vx = best glide speed corresponding to some original gross weight, Wi H2=best glide speed corresponding to some new gross weight, W%

As a result of this relationship, a 10 percent increase in gross weight would require a 5 per­cent increase in glide speed to maintain (L/Z>)mea.- While small variations in gross weight may produce a measurable change in best glide speed, the airplane can tolerate small deviations from the optimum Cu without signif­icant change in(L/D) and glide ratio. For this reason, a standard, single value of glide speed may be specified for a small range ol gross weights at which glide performance can be of importance. A gross weight which is con­siderably different from the normal range will require a modification of best glide speed to maintain the maximum glide ratio.

The effect of altitude on glide performance is insignificant if there is no change in (LjD’)max. Generally, the glide performance of the major­ity of airplanes is subsonic and there is no noticeable variation of with altitude.

Any specific airplane configuration at a partic­ular gross weight will require a specific value of dynamic pressure to sustain flight at the Q, for (Z/D)TOoJ,. Thus, the airplane will have a best glide speed which is a specific value of equivalent airspeed (EAS) independent of altitude. For convenience and simplicity, this best glide speed is specified as a specific value of indicated airspeed (lAS^} and compressibility and position errors are neglected. The prin­cipal effect of altitude is that at high altitude the true airspeed (TAS) and rate of descent along the optimum glide path are increased above the low altitude conditions. However, if (L/D)maj is maintained, the glide angle and glide ratio are identical to the low altitude conditions.

The effect of configuration has been noted pre­viously in that the addition of parasite drag by flaps, landing gear, speed brakes, external stores, etc. will reduce the maximum lift-drag ratio and cause a reduction of glide ratio. In the case where glide distance is of great im­portance, the airplane must be maintained in the clean configuration and flown at (L/D") .

The effect of wind on gliding performance is similar to the effect of wind on cruising range. That is, a headwind will always reduce the glide range and a tailwind will always increase the glide range. The maximum glide range of the airplane in still air will be obtained by flight at (L(D’)1Ma – However, when a wind is present, the optimum gliding conditions may not be accomplished by operation at (L/D’)max. For example, when a headwind is present, the optimum glide speed will be increased to obtain a maximum proportion of ground dis­tance to altitude. In this sense, the increased glide speed helps to minimize the detrimental effect of the headwind. In the case of a tail­wind, the optimum glide speed will be reduced to maximize the benefit of the tailwind. For ordinary wind conditions, maintaining the glide speed best for zero wind conditions will suffice and the loss or gain in glide distance must be accepted. However, when the wind conditions are extreme and the wind velocity is large in comparison with the glide speed, e. g., wind velocity greater than 25 percent of the glide speed, changes in the glide speed must be made to obtain maximum possible ground distance.

THE FLAMEOUT PATTERN. In the case of failure of the powerplant, every effort should be made to establish a well-planned, stabilized approach if a suitable landing area is available. Generally a 360° overhead ap­proach is specified with the approach begin­ning from the ‘ ‘high key’ ’ point of the flameout pattern. The function of a standardized pattern is to provide a flight path well within the capabilities of the airplane and the abilities of the pilot to judge and control the flight path. The flight handbook will generally specify the particulars of the flameout pattern such as the altitude at the high key, glide speeds, use of flaps, etc. Of course, the par­ticulars of the flameout pattern will be de­termined by the aerodynamic characteristics of the airplane. A principal factor is the effect of glide ratio, or (L/D)mai, on the alti­tude required at the high key point at the beginning of the flameout pattern. The air­plane with a low value of (L/D) will require a high altitude at the high key point.

The most favorable situation during a flameout would be for the airplane to in posi­tion to arrive over the intended landing area the altitude for the high key point. In this case, the standard flameout pattern could be utilized. If the airplane does not have suffi­cient glide range to arrive at the landing area with the altitude for the high key point, it is desirable to fit the approach into the lower portions of the standard flameout ap­proach. If it is not possible to arrive at the intended landing area with sufficient altitude to ‘‘play” the approach, serious considera­tion should be given to ejection while suffi­cient altitude remains. Deviations from a well-planned approach such as the standard flameout pattern may allow gross errors in judgment. A typical error of a non-standard or poorly executed flameout approach is the use of excessive angles of bank in turns to correct the approach. Because of the great increase in induced drag at large angles of bank, excessive rates of descent will be incurred and there will be further deviations from a desirable flight path.

The power-off gliding characteristics of the airplane can be simulated in power on flight by certain combinations of engine power setting and position of the speed brake or dive flap. This will allow the pilot to become familiar with the power-off glide performance and the flameout landing pattern. In addition, the simulated flameout pattern is useful during a precautionary landing when the powerplant is malfunctioning and there is the possibility of an actual flameout.

The final approach and landing flare will be particularly critical for the airplane which has’ a low glide ratio but a high best glide speed. These airplane characteristics are typical of the modern configuration of airplane which has

low aspect ratio, sweepback, and high wing loading. Since these airplane characteristics also produce marginal flare capability in power – off flight, great care should be taken to follow the procedure recommended for the specific airplane.

As an example of the power-off glide per­formance of an airplane with low aspect ratio, sweepback, and high wing loading, a best glide speed of 220 knots and a glide ratio of 6 may be typical. In such a case, the rate of descent during the glide at low altitude would be on the order of 3,700 FPM. Any deviations from the recommended landing technique can­not be tolerated because of the possibility of an excessive rate of descent. Either premature flare or delayed flare may allow the airplane to touch down at a rate of descent which would cause structural failure. Because of the mar­ginal flare characteristics in power-off flight, the best glide speed recommended for the land­ing configuration may be well above the speed corresponding to the exact maximum lift-drag ratio. The greater speed reduces induced drag and provides a greater margin for a successful power-off landing flare.

In the extreme case, the power-off glide and landing flare characteristics may be very criti­cal for certain airplane configurations. Thus, a well-planned standard flameout pattern and precise flying technique are necessary and, if very suitable conditions are not available, the recommended alternative is simple: eject!

The variation of wind velocity and direction throughout the atmosphere is important be­cause of its effect on the aerodynamic forces and moments on an airplane. As the airplane traverses this variation of wind velocity and direction during flight, the changes in airflow direction and velocity create changes in the aerodynamic forces and moments and produce a response of the airplane. The variation of airflow velocity along a given direction exists with shear parallel to the flow direction. Hence, the velocity gradients are often re­ferred to as the wind “shear.”

The effect of the vertical gust has important effects on the airplane at high speed because of the possibility of damaging flight loads. The mechanism of vertical gust is illustrated in figure 6.5 where the vertical gust velocity is added vectorially to the flight velocity to produce some resultant velocity. The principal effect of the vertical gust is to produce a change in airplane angle of attack, e. g., a positive (up) gust causes an increase in angle of attack while a negative (down) gust causes a de­crease in angle of attack. Of course, a change in angle of attack will effect a change in lift and, if some critical combination of high gust intensity and high flight speed is encountered, the change in lift may be large enough to cause structural damage.

At low flight speeds during approach, land­ing, and takeoff, the effect of the vertical gust is due to the same mechanism of the change in angle of attack. However, at these low flight speeds, the problem is one of possible incipient stalling and sinking rather than overstress. When the airplane is at high angle of attack, a further increase in angle of attack due to a gust may exceed the critical angle of attack and cause an incipient stalling of the airplane. Also, a decrease in angle of attack due to a gust will cause a loss of lift and allow the airplane to sink. For this reason, any deficiency of airspeed will be quite critical when operating in gusty conditions.

The effect of the horizontal gust differs from the effect of the vertical gust in that the im­mediate effect is a change of airspeed rather than a change in angle of attack. In this sense, the horizontal gust is of little conse­quence in the major airplane airloads and strength limitations. Of greater significance is the response of the airplane to horizontal gusts and wind shear when operating at low flight speeds. The possible conditions in which an airplane may encounter horizontal gusts and wind shear are illustrated in figure 6.5. As the airplane traverses a shear of wind direction, a change in headwind component will exist. Also, a climbing or descending airplane may traverse a shear of wind velocity,

i. e., a wind profile in which the wind velocity varies with altitude.

The response of an airplane is much de­pendent upon the airplane characteristics but certain basic effects are common to all air­planes. Suppose that an airplane is estab­lished in steady, level flight with lift equal to weight, thrust equal to drag, and trimmed so

there is no unbalance of pitching, yawing, or rolling moment. If the airplane traverses a sharp wind shear equivalent to a horizontal gust, the resulting change in airspeed will disturb such an equilibrium. For example, if the airplane encounters a sharp horizontal gust which reduces the airspeed 20percent, the new airspeed (80 percent of the original value} produces lift and drag at the same angle of attack which are 64 percent of the original value. The change in these aerodynamic forces would cause the airplane to accelerate in the direction of resultant unbalance of force. That is, the airplane would accelerate down and forward until a new equilibrium is achieved. In addition, there would be a change in pitching moment which would produce a response of the airplane in pitch.

The response of the airplane to a horizontal gust will differ according to the gust gradient and airplane characteristics. Generally, if the airplane encounters a sharp wind shear which reduces the airspeed, the airplane tends to sink and incur a loss of altitude before equilibrium conditions are achieved. Similarly, if the airplane encounters a sharp wind shear which increases the airspeed, the airplane tends to float and incur a gain of altitude before equilib­rium conditions are achieved.

Significant vertical and horizontal gusts may be due to the terrain or atmospheric conditions. The proximity’of an unstable front or thunder­storm activity in the vicinity of the airfield is likely to create significant wind shear and gust activity at low altitude. During gusty condi­tions every effort must be made for precise con­trol of airspeed and flight path and any changes due to gusts must be corrected by proper con­trol action. Under extreme gusts conditions, it may be advisable to utilize approach, land­ing, and takeoff speeds slightly greater than normal to provide margin for adequate control.