Category AVIATORS

As in the case of landing, the specific tech­niques necessary may vary greatly between various types of airplanes and various oper­ations but certain fundamental principles will be common to all airplanes and all operations. The specific procedures recommended for each airplane type must be followed exactly to insure a consistent, safe takeoff flying tech­nique.

TAKEOFF SPEED AND DISTANCE. The takeoff speed of any airplane is some mini­mum practical airspeed which allows sufficient margin above stall and provides satisfactory control and initial rate of climb.. Depending on the airplane characteristics, the takeoff speed will be some value 5 to 25 percent above the stall or minimum control speed. As such, the takeoff will be accomplished at a certain value of lift coefficient and angle of attack specific to each airplane configuration. As a result, the takeoff airspeed (EAS or CAS’) of any specific airplane configuration is a function of the gross weight at takeoff. Too low an airspeed at takeoff may cause stall, lack of adequate control, or poor initial climb per­formance. An excess of speed at takeoff may provide better control and initial rate of climb but the higher speed requires additional dis­tance and may provide critical conditions for the tires.

The takeoff distance of an airplane is affected by many different factors other than technique and, prior to takeoff, the takeoff distance must be determined and compared with the runway length available. The principal factors affecting the takeoff distance are as follows:

(1) The gross weight of the airplane has a considerable effect on takeoff distance be­cause it affects both takeoff speed and ac­celeration during takeoff roll.

(2) The surface winds must be considered because of the powerful effect of a headwind or tailwind on the takeoff distance. In the case of the cross wind, the component of wind along the runway will be the effective headwind or tailwind velocity. In addi­tion, the component of wind across the run­way will define certain requirements of lateral control power and the limiting compo­nent wind must not be exceeded.

(3) Pressure altitude and temperature can cause a large effect on takeoff distance, es­pecially in the case of the turbine powered airplane. Density altitude will determine the true airspeed at takeoff and can affect the takeoff acceleration by altering the powerplant thrust. The effect of tempera­ture alone is important in the case of the turbine powered aircraft since inlet air tem­perature will affect powerplant thrust. It should be noted that a typical turbojet air­plane may be approximately twice as sensi­tive to density altitude and five to ten times as sensitive to temperature as a representa­tive reciprocating engine powered airplane.

(4) Specific humidity must be accounted for in the case of the reciprocating engine powered airplane. A high water vapor content in the air will cause a definite reduc­tion in takeoff power and takeoff acceler­ation.

(5) The runway condition will deserve con­sideration when the takeoff acceleration is basically low. The runway slope must be compared carefully with the surface winds because ordinary values of runway slope will usually favor choice of the runway with headwind and upslope rather than down – slope and tailwind. The surface condition of the runway has little bearing on takeoff distance as long as the runway is a hard surface.

Each, of these factors must be accounted for and the takeoff distance properly com­puted for the existing conditions. Since obstacle clearance distance is generally a function of the same factors which affect takeoff distance, the obstacle clearance dis­tance is usually related as some proportion of the takeoff distance. Of course, the take­off and obstacle clearance distances related by the handbook data will be obtained by the techniques and procedures outlined In the handbook.

TYPICAL ERRORS. The takeoff distance of an airplane should be computed for each takeoff. A most inexcusable error would be to attempt takeoff from a runway of insufficient length. Familiarity with the airplane hand­book performance data and proper accounting of weight, wind, altitude, temperature, etc., are necessary parts of flying. Conditions of high gross weight, high pressure altitude and temperature, and unfavorable winds create the extreme requirements of runway length, espe­cially for the turbine powered airplane. Under these conditions, use of the handbook data is mandatory and no guesswork can be tolerated.

One typical, error of takeoff technique is the premature or excess pitch rotation of the air­plane. Premature or excess fitch rotation of the airplane may seriously reduce the takeoff accel­eration and increase the takeoff distance. In addition, when the airplane is placed at an excessive angle of attack during takeoff, the airplane may become airborne at too low a speed and the result may be a stall, lack of ade­quate control (especially in a crosswind), or poor initial climb performance. In fact, there are certain low aspect ratio configurations of airplanes which, at an excessive angle of at­tack, will not fly out of ground effect. Thus, over-rotation of the airplane during takeoff may hinder takeoff acceleration or the. initial climb. It is quite typical for an airplane to be placed at an excess angle of attack and become airborne prematurely then settle back to the runway. When the proper angle of attack is assumed, the airplane simply accelerates to the takeoff speed and becomes airborne with suf­ficient initial rate of climb. In this sense, the appropriate rotation and takeoff speeds or an angle of attack indicator must be used.

If the airplane is subject to a sudden pull-up or steep turn after becoming airborne, the result may be a stall, spin, or reduction in initial rate of climb. The increased angle of attack may exceed the critical angle of attack or the in­crease in induced drag may be quite large. For this reason, any clearing turns made immedi­ately after takeoff or deck launch must be slight and well within the capabilities of the air­plane.

In order to obviate some of the problems of a deficiency of airspeed at takeoff, usual result can be an excess of airspeed at takeoff. The principal effect of an excess takeoff airspeed is the greater takeoff distance which results. The general effect is that each 1 percent excess takeoff velocity incurs approximately 2 per­cent additional takeoff distance. Thus, excess speed must be compared with the additional runway required to produce the higher speed. In addition, the aircraft tires may be subject to critical loads when the airplane is at very high rolling speeds and speeds in excess of a basically high takeoff speed may produce damage or failure of the tires.

As with the conditions of landing, excess velocity or deficiency of velocity at takeoff is undesirable. The proper takeoff speeds and angle of attack must be utilized to assure satisfactory takeoff performance.

. The specific techniques of landing flare and touchdown will vary considerably between various types of airplanes. In fact, for certain types of airplanes, a flare from a properly executed approach may not be de­sirable because of the possibility of certain critical dynamic landing loads or because of the necessity for a certain standard of tech­nique when aerodynamic flare characteristics are critical. The landing speed should be the lowest practical speed above the stall or mini­mum control speed to reduce landing distances and arresting loads. Generally, the landing speed will be from 5 to 25 percent above the stall speed depending on the airplane type and the particular operation.

The technique required for the landing will be determined in great part by the aerodynamic characteristics of the airplane. If the airplane characteristics are low wing loading, high L/D, and relatively high lift curve slope, the airplane usually will have good landing flare charac­teristics. If the airplane characteristics are high wing loading, low LID, and relatively low lift curve slope, the airplane may not possess desirable flare characteristics and landing tech­nique may require a minimum of flare to touchdown. These extremes are illustrated by the lift curves of figure 6.4.

In preparation for the landing, several factors must be accounted for because of their effect on landing distance, landing loads, and arrest­ing loads. These factors are:

(1) Landing gross weight must be con­sidered because of its effect on landing speed and landing loads. Since the landing is accomplished at a specific angle of attack or margin above the stall speed, gross weight will define the landing speed. In addition, the gross weight is an important factor in determining the landing distance and energy dissipating requirements of the brakes. There will be a maximum design landing weight specified for each airplane and this limitation must be respected because of critical landing loads, arresting loads, or brake requirements. Of course, any air­plane will have a limiting touchdown rate of descent specified with the maximum land­ing weight and the principal landing load limitations will be defined by the combina­tion. of gross weight and rate of descent at touchdown.

(2) The surface winds must be considered because of the large effect of a headwind or tailwind on the landing distance. In the case of the crosswind, the component of wind along the runway will be the effective headwind or tailwind velocity. Also, the crosswind component across the runway will define certain requirements of lateral control power. The airplane which exhibits large dihedral effect at high lift coefficients is quite sensitive to crosswind and a limiting crosswind component will be defined for the configuration.

(3) Pressure altitude and temperature will affect the landing distance because of the effect on the true airspeed for landing. Thus, pressure altitude and temperature must be considered to define the density altitude.

(4) The runway condition must be con­sidered for its effect on lauding distances. Runway slope of ordinary values will ordi­narily favor selection of a runway for a favorable headwind at landing. The surface condition of the runway will determine braking effectiveness and ice or water on the runway may produce a considerable increase in the minimum landing distance.

Thus, preparation for the landing must in­clude determination of the landing distance of the airplane and comparison with the runway length available. Use of the angle of attack indicator and the mirror landing system will assist the pilot in effecting touchdown at the desired location with the proper airspeed. Of course, the landing is not completed until the airplane is slowed to turn off the runway. Control of the airplane must be maintained after the touchdown and proper technique must be used to decelerate the airplane.

TYPICAL ERRORS. There are many un­desirable consequences when basic principles and specific procedures are not followed during the approach and landing. Some of the typical errors involved in landing accidents are out­lined in the following discussion.

The steep, low power approach leads to an excessive rate of descent and the possibility of a hard landing. This is particularly the case for the modern, low aspect ratio, swept wing airplane configuration which incurs very large induced drag at low speeds and does not have very conventional flare characteristics. For this type of airplane in a steep, low power approach, an increased angle of attack without a change of power setting may not cause a reduction of rate of descent and may even in­crease the rate of descent at touchdown. For this reason, a moderate stabilized approach is necessary and the principal changes in rate of descent must be controlled by changes in power setting and principal changes in airspeed must be controlled by changes in angle of attack.

’An excessive angle of attack during the ap­proach and landing implies that the airplane is being operated at too low an airspeed. Of course, excessive angle of attack may cause the airplane to stall or spin and the low altitude may preclude recovery. Also, the low aspect ratio configuration at an excessively low air­speed will incur very high induced drag and will necessitate a high power setting or other­wise incur an excessive rate of descent. An additional problem is created by an excessive angle of attack for the airplane which exhibits a large dihedral effect at high lift coefficients. In this case, the airplane would be more sensi­tive to crosswinds and adequate lateral control may not be available to effect a safe landing at a critical value of crosswind.

Excess airspeed at landing is just as undesira­ble as a deficiency of airspeed. An excessive airspeed at landing will produce an undesirable increase in landing distance and the energy to be dissipated by the brakes for the field landing or excessive arresting loads for the shipboard landing. In addition, the excess airspeed is a corollary of too low an angle of attack and the airplane may contact the deck or runway nose wheel first and cause damage to the nose wheel or begin a porpoising of the airplane. During a flare to landing, any excess speed will be difficult to dissipate due to the reduction of drag due to ground effect. Thus, if the air­plane is held off with excess airspeed the air­plane will “float” with the consequence of a barrier engagement, barricade engagement, bolter, or considerable runway distance used before touchdown.

A fundamental requirement for a good land­ing is a well planned and executed approach. The possibility of errors during the landing process is minimized when the airplane is brought to the point of touchdown with the proper glide path and airspeed. With the proper approach, there is no need for drastic changes in the flight path, angle of attack, or power setting to accomplish touchdown at the intended point on the deck or runway. Late corrections to line up with the deck or diving for the deck are common errors which eventu­ally result in landing accidents. Accurate control of airspeed and glide path are ab­solutely necessary and the LSO, angle of attack indicator, and the mirror landing system pro­vide great assistance in accurate control of the airplane.

The specific techniques necessary during the phase of approach and landing may vary con­siderably between various types of airplanes and various operations. However, regardless of the airplane type or operation, there are certain fundamental principles which will de­fine the basic techniques of flying during ap­proach and landing. The specific procedures recommended for each airplane type must be followed exactly to insure a consistent, safe landing technique.

THE APPROACH. The approach must be conducted to provide a stabilized, steady flight path to the intended point of touchdown. The approach speed specified for an airplane must provide sufficient margin above the stall speed or minimum control speed to allow satisfactory control and adequate maneuverability. On the other hand, the approach speed must not be greatly in excess of the touchdown speed or a large reduction in speed would be necessary prior to ground contact. Generally, the ap­proach speed will be from 10 to 30 percent above the stall speed depending on the air­plane type and the particular operation.

During the approach, the pilot must attempt to maintain a smooth flight path and prepare for the touchdown. A smooth, steady ap­proach to landing will minimize the transient items of the flight path and provide the pilot better opportunity to perceive and orientate the airplane along the desired flight path. Steep turns must be avoided at the low speeds of the approach because of the increase in drag and stall speed in the turn. Figure 6.4 illus­trates the typical change in thrust required caused by a steep turn. A steep turn may cause the airplane to stall or the large increase in in­duced drag may create an excessive rate of descent. In either case, there may not be suf­ficient altitude to effect recovery. If the air­plane is not properly lined up on the final ap­proach, it is certainly preferable to take a waveoff and go around rather than “press on regardless” and attempt to salvage a decent landing from a poor approach.

The proper coordination of the controls is an absolute necessity during the approach. In this sense, due respect must be given to the primary control of airspeed and rate of descent for the conditions of the steady approach. Thus, the proper angle of attack will produce the desired approach airspeed; too low an angle of attack will incur an excess speed while an excessive angle of attack will produce a deficiency of speed and may cause stall or con­trol problems. Once the proper airspeed and angle of attack are attained the primary control of rate of descent during the steady approach will be the power setting. For example, if it is realized that the airplane is above the de­sired glide path, a more nose-down attitude without a decrease in power setting will result in a gain in airspeed. On the other hand, if it is realized that the airplane is below the desired glide path, a more nose-up attitude without an increase in power setting will simply allow the airplane to fly more slowly and—in the region of reversed command—eventually produce a greater rate of descent. For the conditions of steady flight, angle of attack is the primary control of airspeed and power setting is the primary control of rate of climb and descent. This is especially true during the steady ap­proach to landing. Of course, the ability of the powerplant to produce rapid changes in thrust will affect the specific technique to be used. If the powerplant is not capable of pro­ducing immediate controlled changes in thrust, the operating technique must1 account for this

deficiency. It is most desirable that the power – plant be capable of effecting rapid changes in thrust to allow precise control of the airplane during approach.

The type of approach path is an important factor since it affects the requirement of the flare, the touchdown rate of descent, and—to some extent—the ability to control the point of touchdown. Approach path A of figure 6.4 depicts the steep, low power approach. Such a flight path generally involves a low power setting near idle conditions and a high rate of descent. Precise control of the air­plane is difficult and an excess airspeed usually results from an approach path similar to A. Waveoff may be difficult because of the re­quired engine acceleration and the high rate of descent. In addition, the steep approach path with high rate of descent requires con­siderable flare to reduce the rate of descent at touchdown. This extreme flare requirement will be difficult to execute with consistency and will generally result in great variation in the speed, rate of descent, and point of touchdown.

Approach path C of figure 6.4 typifies the long, shallow approach with too small an inclination of the flight path.- Such a flight path requires a relatively high power setting and a deficiency of airspeed is a usual conse­quence. This extreme of an approach path is not desirable because it is difficult to control the point of touchdown and the low speed may allow the airplane to settle prematurely short of the intended landing touchdown.

Some approach path between the extremes of A and C must be selected, e. g., flight path B. The desirable approach path must not incur excessive speed and rate of descent or require excessive flaring prior to touchdown. Also, some moderate power setting must be required which will allow accurate control of the flight path and provide suitable waveoff characteristics. The approach flight path cannot be too shallow for excessive power setting may be required and it may be difficult to judge and control the point of touchdown. The LSO, mirror landing system, and various approach lighting systems will aid the pilot in achieving the desired approach flight path.

The usual errors during the takeoff and landing phases of flight involve improper con­trol of airspeed and altitude along some desired flight path. Any errors of technique are ampli­fied when an adequate visual reference is not available to the pilot. It is necessary to provide the pilot with as complete as possible visual reference field to minimize or eliminate any errors in perception and orientation. The angle of attack indicator and the mirror land­ing system assist the pilot during the phases of takeoff and landing and allow more consistent, precise control of the airplane.

THE ANGLE OF ATTACK INDICATOR. Many specific aerodynamic conditions exist at particular angles of attack for the airplane. Generally, the conditions of stall, landing ap­proach, takeoff, range, endurance, etc., all occur at specific values of lift coefficient and specific airplane angles of attack. Thus, an instrument to indicate or relate airplane angle of attack would be a valuable reference to aid the pilot.

When the airplane is at high angles of attack it becomes difficult to provide accurate indica­tion of airspeed because of the possibility of large position errors. In fact, for low aspect ratio airplane configurations at high angles of attack, it is possible to provide indications of angle of attack which are more accurate than indications of airspeed. As a result, an angle of attack indicator can be of greatest utility at the high angles of attack.

A particular advantage of an angle of attack indicator is that the indicator is not directly affected by gross weight, bank angle, load factor, velocity, or density altitude. The typical lift curve of figure 6.3 illustrates the variation of lift coefficient, CL, with angle of attack a. When a particular aerodynamic configuration is in subsonic flight, each angle of attack produces a particular value of lift coefficient. Of course, a point of special interest on the lift curve is the maximum lift coefficient, CLmax. Angles of attack greater than that for CLjaax produce a decrease in lift coefficient and constitute the stalled condition of flight. Since CLjnax occurs at a particular angle of attack, any device to provide a stall warning should be predicated on the function of this critical angle o( attack. Under these conditions, stall of the airplane may take place at various airspeeds depending on gross weight, load factor, etc., but always the same angle of attack.

In order to reduce takeoff and landing dis­tances and minimize arresting loads, takeoff and landing will be accomplished at minimum practical speeds. The takeoff and landing speeds must provide sufficient margin above the stall speed (or minimum control speed) and are usually specified at some fixed per­centages of the stall speed. As such, takeoff, approach, and landing will be accomplished at specific values of lift coefficient and, thus, particular angles of attack. For example, assume that point A on the lift curve is defined as the proper aerodynamic condition for the landing approach. This condition exists as a particular lift coefficient and angle of attack for a specific aerodynamic configuration. When the airplane is flown in a steady flight path at the prescribed angle of attack, the resulting airspeed will be appropriate for the airplane gross weight. Any variation in gross weight will simply alter the airspeed necessary to provide sufficient lift. The use of an angle of attack indicator to maintain the recom­mended angle of attack will insure that the airplane is operated at the proper approach speed—not too low or too high an airspeed.

In addition to the use of the angle of attack indicator during approach and landing, the instrument may be used as a principal reference during takeoff. The use of the angle of attack indicator to assume the proper takeoff angle of attack will prevent both over-rotation and excess takeoff speed. Also, the angle of at­tack indicator may be applicable to assist in control of the airplane for conditions of range, endurance, maneuvers, etc.

THE MIRROR LANDING SYSTEM. A well planned, stabilized approach is a funda­mental requirement for a good landing. How­ever, one of the more difficult problems of perception and orientation is the positioning of the airplane along a proper flight path dur­ing approach to landing. While various de­vices are possible, the most successful form of glide path indicator applicable to both field and shipboard operations is the mirror landing system. The function of the mirror landing system is to provide the pilot with an accurate visual reference for a selected flight path which has the desired inclination and point of touch­down. Utilization of the mirror system will allow the pilot to position the airplane along the desired glide path and touch down at the desired point. When the proper glide path inclination is set, the pilot can be assured that the rate of descent will not be excessive and a foundation is established for a successful landing.

The combination of the angle of attack in­dicator and the mirror landing system can provide an excellent reference for a landing technique. The use of the angle of attack indicator will provide the airplane with the proper airspeed while the mirror system refer­ence will provide the desired flight path. When shipboard operations are conducted without the mirror system and angle of attack indicator, the landing signal officer must provide the immediate reference of airspeed and flight path. The LSO must perceive and judge the angle of

attack (and, hence, airspeed) and the flight path of the landing aircraft and signal correc­tions to be made in order to achieve the desired flight path and angle of attack. Because of the field of orientation available to the LSO, he is able to perceive the flight path and angle of attack more accurately than the pilot with­out an angle of attack indicator and mirror landing system.

While the previous chapters have presented the detailed parts of the general field of aero­dynamics, there remain various problems of flying which require the application of princi­ples from many parts of aerodynamics. The application of aerodynamics to these various problems of flying will assist the Naval Aviator in understanding these problems and develop­ing good flying techniques.

PRIMARY CONTROL OF AIRSPEED AND ALTITUDE

For the conditions of steady flight, the air­plane must be in equilibrium. Equilibrium will be achieved when there is no unbalance of force or moment acting on the airplane. If it is assumed that the airplane is trimmed so that no unbalance of pitching, yawing, or rolling moments exists, the principal concern is for

the forces acting on the airplane, i. e., lift, thrust, weight, and drag.

ANGLE OF ATTACK VERSUS AIRSPEED. In order to achieve equilibrium in the vertical direction, the net lift must equal the airplane weight. This is a contingency of steady, level flight or steady climbing and descending flight when the flight path inclination is slight. A refinement of the basic lift equation defines the relationship of speed, weight, lift coefficient, etc., for the condition of lift equal to weight.

or

FE=17.2y^

where

V= velocity, knots ([TAS)

VE = equivalent airspeed, knots (EAS) W= gross weight, lbs.

T=wing surface area, sq. ft.

WjS= wing loading, psf v = altitude density ratio Cb = lift coefficient

From this relationship it is appreciated that a given configuration of airplane with a specific wing loading, W/S, will achieve lift equal to weight at particular combinations of velocity, V, and lift coefficient, CL. In steady flight, each equivalent airspeed demands a particular value of Cb and each value of Cb demands a particular equivalent airspeed to provide lift equal to weight. Figure 6.1 illustrates a typical lift curve for an airplane and shows the relation­ship between CL and a, angle of attack. For this relationship, some specific value of a will create a certain value of CL for any given aero­dynamic configuration.

For the conditions of steady flight with a given airplane, each angle of attack corre­sponds to a specific airspeed. Each angle of attack produces a specific value of CL and each value of Ci requires a specific value of equiva­lent airspeed to provide lift equal to weight. Hence, angle of attack is the primary control of airspeed in steady flight, If an airplane is es­tablished in steady, level flight at a particular airspeed, any increase in angle of attack will result in some reduced airspeed common to the increased CL. A decrease in angle of attack will result in some increased airspeed com­mon to the decreased CL. As a result of the change in airspeed, the airplane may climb or descend if there is no change in power setting but the change in airspeed was provided by the change in angle of attack. The state of the airplane during the change in speed will be some transient condition between the original and final steady state conditions.

Primary control of airspeed in steady flight by angle of attack is an important principle. With some configurations of airplanes, low speed flight will bring about a low level of longitudinal stick force stability and possi­bility of low airplane static longitudi­nal stability. In such a case, the “feel” for airspeed will be light and may not furnish a ready reference for easy control of the air­plane. In addition, the high angles of attack common to low speed flight are likely to pro­vide large position errors to the airspeed indi­cating system. Thus, proper control of air­speed will be enhanced by good “attitude” flying or—when the visual reference field is poor—an angle of attack indicator.

RATE OF CLIMB AND DESCENT. In order for an airplane to achieve equilibrium at constant altitude, lift must be equal to weight and thrust must be equal to drag. Steady, level flight requires equilibrium in both the vertical and horizontal directions. For the case of climbing or descending flight condi­tions, a component of weight is inclined along the flight path direction and equilibrium is achieved when thrust is not equal to the drag. When the airplane is in a steady climb or descent, the rate of climb is related by the following expression:

fPa-Pr

ЖС^-ЪЪ&Ц^рг-)

where

RC=rate of climb, ft. per min.

Pa = propulsive power available; h, p.

Pr=power required for level flight, h. p.

W= gross weight, lbs.

From this relationship it is appreciated that the rate of climb in steady flight is a direct function of the difference between power avail­able and power required. If a given airplane configuration is in lift-equal-to-weight flight at some specific airspeed and altitude, there is a specific power required to maintain these conditions. If the power available from the powerplant is adjusted to equal the power required, the rate of climb is zero (Pa—Pr=0′). This is illustrated in figure 6.1 where the power available is set equal to the power required at velocity (A). If the airplane were in steady level flight at velocity (A), an increase in power available would create an excess of power which will cause a rate of climb. Of course, if the speed were allowed to increase by a decreased angle of attack, the increased power setting could simply maintain altitude at some higher airspeed. However, if the original aerodynamic conditions are maintain­ed, speed is maintained at (Л) and an increased power available results in a rate of climb. Also, a decrease in power available at point (A) will produce a deficiency in power and result in a negative rate of climb (or a rate of descent). For this reason, it is apparent that power setting is the primary control of altitude in steady flight. There is the direct correlation between the excess power ([Pa—Pr), and the airplane rate of climb, RC.

FLYING TECHNIQUE. Since the condi­tions of steady flight predominate during a majority of all flying, the fundamentals of flying technique are the principles of steady flight:

(1) Angle of attack is the primary control of airspeed.

(2) Power setting is the primary control

of altitude, i. e., rate of clirhb/descent.

With the exception of the transient conditions of flight which occur during maneuvers and acrobatics, the conditions of steady flight will be applicable during such steady flight condi­tions as cruise, climb, descent, takeoff, ap­proach, landing, etc. A clear understanding of these two principles will develop good, safe flying techniques applicable to any sort of airplane.

The primary control of airspeed during steady flight conditions is the angle of attack. However, changes in airspeed will necessitate changes in power setting to maintain altitude because of the variation of power required with velocity. The primary control of altitude (rate of climb/descent) is the power setting. If an airplane is being flown at a particular airspeed in level flight, an increase or decrease in power setting will result in a rate of climb or descent at this airspeed. While the angle of attack must be maintained to hold airspeed in steady flight, a change in power setting, will necessitate a change in attitudeto. accommodate the new flight path direction. These princi­ples form the basis for "attitude” flying tech­nique, i. e., "attitude plus, power equals per­formance," and provide, a background for good instrument flying technique as well as good flying technique for all ordinary flying conditions.

One of the most important phases of flight is the landing approach and it is during this phase of flight that the principles of steady flight are so applicable. If, during the landing approach, it is realized that ;the airplane is below the desired glide path, an increase in nose up attitude will not insure that the airplane will climb to the desired glide path. In fact, an increase in nose-up attitude may produce a greater rate of descent and cause the airplane to sink more below the desired glide path. At a given airspeed, only an increase in power setting can cause a rate of climb (or lower rate of descent) and an in­

crease in nose up attitude without the appro­priate power change only controls the airplane to a lower speed.

REGION OF REVERSED COMMAND

The variation of power or thrust required with velocity defines the power settings neces­sary to maintain steady level flight at various airspeeds. To simplify the situation, a gener­ality could be assumed that the airplane con­figuration and. altitude define a variation of power setting required (jet thrust required or prop power required) versus velocity. This general variation of required power setting versus velocity is illustrated by the first graph of figure 6.2. This curve illustrates the fact that at low speeds near the stall or minimum control speed the power setting required for steady level flight is quite high. However, at low speeds, am increase in speed reduces the required power setting until some minimum value is reached at the conditions for maximum endurance. Increased speed beyond the con­ditions for maximum endurance will then increase the power setting required for steady level flight.

REGIONS OF NORMAL AND REVERSED COMMAND. This typical variation of re­quired power setting with speed allows a sort of terminology to be assigned to specific regimes of velocity. Speeds greater than the speed for maximum endurance require increas­ingly greater power settings to achieve steady, level flight. Since the normal command of flight assumes a higher power setting will achieve a greater speed, the regime of flight speeds greater than the speed for minimum required power setting is termed the “region of normal command.” Obviously, parasite drag or parasite power predominates in this regime to produce the increased power setting required with increased velocity. Of course, the major items of airplane flight performance take place in the region of normal command.

Flight speeds below the speed for maximum endurance produce required power settings

which increase with a decrease in speed. Since the increase in required power setting with decreased velocity is contrary to the normal command of flight, the regime of flight speeds between the speed for minimum required power setting and the stall speed (or minimum control speed) is termed the “region of re­versed command.” In this regime of flight, a decrease in airspeed must be accompanied by an increased power setting in order to main­tain steady flight. Obviously, induced drag or induced power required predominates in this regime to produce the increased power setting required with decreased velocity. One fact should be made clear about the region of reversed command: flight in the “reversed" region of command does not imply that a decreased power setting will bring about a higher airspeed or an increased power setting will produce a lower airspeed. To be sure, the primary control of airspeed is not the power setting. Flight in the region of re­versed command only implies that a higher airspeed will require a lower power setting and a lower airspeed will require a higher power setting to hold altitude.

Because of the variation of required power setting throughout the range of flight speeds, it is possible that one particular power setting may be capable of achieving steady, level flight at two different airspeeds. As shown on the first curve of figure 6.2, one given power setting would meet the power requirements and allow steady, level flight at both points I and 2. At speeds lower than point 2, a deficiency of power | would exist and a rate of descent would be in­curred. Similarly, at speeds greater than point 1, a deficiency of power would exist and the I airplane would descend. The speed range be­tween points 1 and 2 would provide an excess of power and climbing flight would be pro­duced.

FEATURES OF FLIGHT IN THE NOR­MAL AND REVERSED REGIONS OF COM­MAND. The majority of all airplane flight is conducted in the region of normal command,

e. g., cruise, climb, maneuvers, etc. The region of reversed command is encountered primarily in the low speed phases of flight during takeoff and landing. Because of the extensive low speed flight during carrier operations, the Naval Aviator will be more familiar with the region of reversed command than the ordinary pilot.

The characteristics of flight in the region’of normal command are illustrated at point A on the second curve of figure 6.2. If the airplane is established in steady, level flight at point A, lift is equal to weight and the power available is set equal to the power required. When the airplane is disturbed to some airspeed slightly greater than point A, a power deficiency exists and, when the airplane is disturbed to some air­speed slightly lower than point A, a power excess exists. This relationship provides a tendency for the airplane to return to the equili­brium of point A and resume the original flight condition following a disturbance. Also, the static longitudinal stability of the airplane tends to return the airplane to the original trimmed CL and velocity corresponding to this CL. The phugoid usually has most satisfactory qualities at low values of CL so the high speed of the region of normal command provides little tendency of the airplane’s airspeed to vary or wander about.

With all factors considered, flight in the region of normal’ command is characterized by a relatively strong tendency of the airplane to maintain the trim speed quite naturally. How­ever, flight in the region of normal command can lead to some unusual and erroneous impres­sions regarding proper flying technique. For example, if the airplane is established at point A in steady level flight, a controlled increase in airspeed without a change in power setting will create a deficiency of power and cause the airplane to descend. Similarly, a controlled decrease in airspeed without a change in power setting will create an excess of power and cause the airplane to climb. This fact, coupled with the transient motion of the airplane when the angle of attack is changed rapidly, may lead to the impression that rate of climb and descent can be controlled by changes in angle of attack. While such is true in the region of normal com­mand, for the conditions of steady flight, pri­mary control of altitude remains the power setting and the primary control of airspeed re­mains the angle of attack. The impressions and habits that can be developed in the region of normal command can bring about disastrous consequences in the region of reversed com­mand.

The characteristics of flight in the region of reversed command are illustrated at point В on the second curve of figure 6.2. If the air­plane is established in steady, level flight at point B, lift is equal to weight and the power available is set equal to the power required. When the airplane is disturbed to some air­speed slightly greater than point B, an excess of power exists and, when the airplane is dis­turbed to some airspeed slightly lower than point B, a deficiency of power exists. This relationship is basically unstable because the variation of excess power to either side of point В tends to magnify any original dis­turbance. While the static longitudinal sta­bility of the airplane tends to maintain the original trimmed CL and airspeed correspond­ing to that CL, the phugoid usually has the least satisfactory qualities at the high values of CL corresponding to low speed flight.

When all factors are considered, flight in the region of reversed command is characterized by a relatively weak tendency of the airplane to maintain the trim speed naturally. In fact it is likely that the airplane will exhibit no inherent tendency to maintain the trim speed in this regime of flight. For this reason, the pilot must give particular attention to precise control of airspeed when operating in the low flight speeds of the region of reversed command.

While flight in the region of normal com­mand may create doubt as to the primary con­trol of airspeed and altitude, operation in the region of reversed command should leave little

doubt about proper flying techniques. For example, if the airplane is established at point В in level flight, a controlled increase in air­speed (by reducing angle of attack) without change in power setting will create an excess of power at the higher airspeed and cause the airplane to climb. Also, a controlled decrease in airspeed (by increasing angle of attack) without a change of power setting will create a deficiency of power at the lower airspeed and cause the airplane to descend. This rela­tionship should leave little doubt as to the primary control of airspeed and altitude.

The transient conditions during the changes in airspeed in the region of reversed command are of Interest from the standpoint of landing flare characteristics. Suppose the airplane is in steady flight at point В and the airplane angle of attack is increased to correspond with the value for the lower airspeed of point C (see fig. 6.2). The airplane would not instanta­neously develop the lower speed and rate of descent common to point C but would approach the conditions of point C through some tran­sient process depending on the airplane char­acteristics. If the airplane characteristics are low wing loading, high L/D, and high lift curve slope, the increase in angle of attack at point В will produce a transient motion in which curvature of the flight path demonstrates a definite flare. That is, the increase in angle of attack creates a momentary rate of climb (or reduction of rate of descent) which would be accompanied by a gradual loss of airspeed. Of course, the speed eventually decreases to point C and the steady state rate of descent is achieved. If the airplane characteristics are high wing loading, low L/D, and low lift curve slope, the increase in angle of attack at point В may produce a transient motion in which the airplane does not flare. That is, the increase in angle of attack may produce such rapid re­duction of airspeed and increase in rate of descent that the airplane may be incapable of a flaring flight path without an increase in power setting. Such characteristics may neces­sitate special landing techniques, particularly in the case of a flameout landing.

Operation in the region of reversed command does not imply that great control difficulty and dangerous conditions will exist. However, flight in the region of reversed command does amplify any errors of basic flying technique. Hence, proper flying technique and precise control of the airplane are most necessary in the region of reversed command.

Accumulated periods of overstress can create a very detrimental effect on the useful service life of any structural component. This fact is certain and irreversible. Thus, the opera­tion of the airplane, powerplant, and various systems must be limited to design values to prevent failure or excessive maintenance costs early in the anticipated service life. The operating limitations presented in the hand­book must be adhered to in a very strict fashion.

In many cases of modern aircraft structures it is very difficult to appreciate the effect of a moderate overstress. This feature is due in great part to the inherent strength of the materials used in modern aircraft construction. As a general airframe static strength require­ment, the primary structure must not expe­rience objectionable permanent deformation at limit load or failure at 150 percent of limit load (ultimate load is 1.5 times limit load). To satisfy each part of the requirement, limit load must not exceed the yield stress and ulti­mate load must not exceed the ultimate stress capability of the parts.

Many of the high strength materials used in aircraft construction have stress-strain dia­grams typical of figure 5-6. One feature of these materials is that the yield point is at some stress much greater than two-thirds of the ultimate stress. Thus, the critical design condition is the ultimate load. If 150 percent of limit load corresponds to ultimate stress of the material, 100 percent of limit load corre­sponds to a stress much lower than the yield stress. Because of the inherent properties of the high strength material and the ultimate factor of safety of 1.5, the limit load condition is rarely the critical design point and usually possesses a large positive margin of static strength. This fact alone implies that the structure must be grossly overstressed to pro­duce damage easily visible to the naked eye. This lack of immediate visible damage with “overstress” makes it quite difficult to recog­nize or appreciate the long range effect.

A reference point provided on the stress strain diagram of figure 5.6 is a stress termed

the ‘ ‘endurance limit. ” If the operating cyclic stresses never exceed this “endurance limit" an infinite (or in some cases “near infinite”) num­ber of cycles can be withstood without fatigue failure. No significant fatigue damage accrues from stresses below the endurance limit but the value of this endurance limit is approxi­mately 30 to 50 percent of the yield strength for the light alloys used in aircraft construc­tion. The rate of fatigue damage caused by stresses only slightly above the endurance limit is insignificant. Even stresses near the limit load do not cause a significant accumulation of fatigue damage if the frequency of application is reasonable and within the intended mission requirement. However, stresses above the limit load—and especially stresses well above the limit load—create a very rapid rate of fatigue damage.

A puzzling situation then exists. “Over­stress” is difficult to recognize because of the

inherent high yield strength and low ductility of typical aircraft metals. These same over­stresses cause high rate of fatigue damage and create premature failure of parts in service. The effect of accumulated overstress is the formation and propagation of fatigue cracks. While it is sure that fatigue crack always will be formed before final failure of a part, accumu­lated overstress is most severe and fatigue provoking at the inevitable stress concentra­tions. Hence, disassembly and detailed inspec­tion is both costly and time-consuming. To prevent Іп-service failures of a basically sound structure, the part must be properly maintained and operated within the design “envelope.” Examples of in-service fatigue failures are shown in figure 5.7.

The operation of any aircraft and powerplant must be conducted within the operating limita­tions prescribed in the flight handbook. No hearsay or rumors can be substituted for the

accepted data presented in the aircraft hand­book. All of the various static strength, service life, and aeroelastic effects must be given proper respect. An airplane can be over­stressed with the possibility that no immediate damage is apparent. A powerplant may be operated past the specified time, speed, or temperature limits without immediate appar­ent damage. In each case, the cumulative effect will tell at some later time when in­service failures occur and maintenance costs increase.

The most critical loads on the landing gear occur at high gross weight and high rate of descent at touchdown. Since the landing gear has requirements of static strength and fatigue strength similar to any other com­ponent, overstress must be avoided to prevent failure and derive the anticipated service life from the components.

The most significant function of the landing gear is to absorb the vertical energy of the air­craft at touchdown. An aircraft at a given weight and rate of descent at touchdown has a certain kinetic energy which must be dis­sipated in the shock absorbers of the landing gear. If the energy were not absorbed at touchdown, the aircraft would bounce along similar to an automobile with faulty shock absorbers. As the strut deflects on touchdown, oil is forced through an orifice at high velocity and the energy of the aircraft is absorbed. To have an efficient strut the orifice size must be controlled with a tapered pin to absorb the energy with the most uniform force on the strut.

The vertical landing loads resulting at touch­down can be simplified to an extent by assum­ing the action of the strut to produce a uni­formly accelerated motion of the aircraft. T^ landing load factor for touchdown at a consta rate of descent can be expressed by the follow­ing equation:

n= landing load factor—the ratio of the load in the strut, F, to the weight, W

ROD = rate of descent, ft. per sec.

g= acceleration due to gravity = 32 ft. per sec.2

S= effective stroke of the stmt, ft.

As an example, assume that an aircraft touches down at a constant rate of descent of 18 ft. per sec. and the effective stroke of the strut is 18 inches (1.5 ft ). The landing load factor for the condition would be 3-37; the average force would be 3-37 times the weight of the aircraft. (Note: there is no specific correlation between the landing load factor and the indication of a cockpit mounted flight accelerometer. The response of the instrument, its mounting, and the onset of landing loads usually prevent direct correlation.)

This simplified equation points out two im­portant facts. The effective stroke of the strut should be large to minimize the loads since a greater distance of travel reduces the force necessary to do the work of arresting the ver­tical descent of the aircraft. This should

emphasize the necessity of proper maintenance of the struts. An additional fact illustrated is that the landing load factor varies as the square of the touchdown rate of descent. Therefore, a 20 percent higher rate of descent increases the landing load factor 44 percent. This fact should emphasize the need for proper landing technique to prevent a hard landing and over­stress of the landing gear components and associated structure.

The effect of landing gross weight is two­fold. A higher gross weight at some landing load factor produces a higher force in the landing gear. The higher gross weight re­quires a higher approach speed and, if the same glide path is used, a higher rate of descent results. In addition to the principal vertical loads on the landing gear, there are varied side loads, wheel spin up and spring back loads, etc., all of which tend to be more critical at high gross weight, high touchdown ground speed, and high rate of descent.

The function of the landing gear as a shock absorbing device has an important application when a forced landing must be accomplished on an unprepared surface. If the terrain is rough and the landing gear is not extended, initial contact will be made with relatively solid structure and whatever energy is ab­sorbed will be accompanied by high vertical accelerations. These high vertical accelera­tions encountered with a gear-up landing on an unprepared surface are the source of a very incapacitating type injury—vertical compres­sion fracture of the vertebrae. Unless some peculiarity of the configuration makes it inadvisable, it is generally recommended that the landing gear be down for forced landing on an unprepared surface. (Note: for those prone to forget, it is also recommended that the gear be down for landing on prepared surfaces.)

Many different factors may be of structural importance in high speed flight. Any one or combination of these factors may be encount­ered if the airplane is operated beyond the limit (or redline) airspeed.

At speeds beyond the limit speed the air­plane may encounter a critical gust. This is especially true of a high aspect ratio airplane with a low limit load factor. Of course, this is also an important consideration for an air­plane with a high limit load factor if the gust should be superimposed on a maneuver. Since the gust load factor increment varies directly with airspeed and gust intensity, high airspeeds must be avoided in turbulent conditions.

When it is impossible to avoid turbulent conditions and the airplane must be subject to gusts, the flight condition must be properly controlled to minimize the effect of turbulence. If possible, the airplane airspeed and power should be adjusted prior to entry into turbu­lence to provide a stabilized attitude. Ob­viously, penetration of turbulence should not be accomplished at an excess airspeed because of possible structural damage. On the other hand, an excessively low speed should not be chosen to penetrate turbulence for the gusts may cause stalling of the aircraft and difficulty of control. To select a proper penetration airspeed the speed should not be excessively high or low—the two extremes must be tempered. The “maneuver” speed is an im­portant reference point since it is the highest speed that can be taken to alleviate stall due to gust and the lowest speed at which limit load factor can be developed aer©dynamically. The optimum penetration speed occurs at or very near the maneuver speed.

Aileron reversal is a phenomenon particular to high speed flight. When in flight at very high dynamic pressures, the wing torsional deflections which occur with aileron deflection are considerable and cause noticeable change in aileron effectiveness. The deflection of an aileron on a rigid wing creates a change in lift and produces a rolling moment. In addition the deflection of the control surface creates a twisting moment on the wing. When the actual elastic wing is subject to this condition at high dynamic pressures, the twisting mo­ment produces measurable twisting deforma­tions which affect the rolling performance of the aircraft. Figure 5 5 illustrates this process and the effect of airspeed on aileron effective­ness. At some high dynamic pressure, the

twisting deformation will be great enough to nullify the effect on aileron deflection and the aileron effectiveness will”be zero. Since speeds above this point create rolling moments op­posite to the direction controlled, this point is termed the “aileron reversal speed.” Oper­ation beyond the reversal speed would create an obvious control difficulty. Also, the ex­tremely large twisting moments which produce loss of aileron effectiveness create large twist­ing moments capable of structural damage.

In order to prevent loss of aileron effective­ness at high airspeeds, the wing must have high torsional stiffness. This may be a feature difficult to accomplish in a wing of very thin section and may favor the use of inboard ailer­ons to reduce the twisted span length and effectively increase torsional stiffness. The use of spoilers for lateral control minimizes the twisting moments and alleviates the reversal problem.

Divergence is another phenomenon common to flight at high dynamic pressures. Like aileron reversal, it is an effect due to the inter­action of aerodynamic forces and elastic deflec­tions of the structure. However, it differs from aileron reversal in that it is a violent instability which produces immediate failure. Figure 5-5 illustrates the process of instability. If the surface is above the divergence speed, any disturbance precipitates this sequence. Any change in lift takes place at the aerody­namic center of the section. The change in lift ahead of the elastic axis produces a twist­ing moment and a consequent twisting deflec­tion. The change in angle of attack creates greater lift at the a. c,, greater twisting deflec­tion, more lift, etc., until failure occurs.

At low flight speeds where the dynamic pressure is low, the relationship between aero­dynamic force buildup and torsional deflection is stable. However, the change in lift per angle of attack is proportional to V2 but the structural torsional stiffness of the wing re­mains constant. This relationship implies that at some high speed, the aerodynamic force buildup may overpower the resisting torsional stiffness and “divergence” will occur. The divergence speed of the surfaces must be suf­ficiently high that the airplane does not en­counter this phenomenon within the normal operating envelope. Sweepback, short span, and high taper help raise the divergence speed.

Flutter involves aerodynamic forces, inertia forces and the elastic properties of a surface. The distribution of mass and stiffness in a structure determine certain natural frequencies and modes of vibration. If the structure is sub­ject to a forcing frequency near these natural frequencies, a resonant condition can result with an unstable oscillation. The aircraft is subject to many aerodynamic excitations while in operation and the aerodynamic forces at various speeds have characteristic properties for rate of change of force and moment. The aerodynamic forces may interact with the structure in a fashion which may excite or negatively damp the natural modes of the structure and allow flutter. Flutter must not occur within the normal flight operating en­velope and the natural modes must be damped if possible or designed to occur beyond the limit speed. A Typical flutter mode is illus­trated in figure 5-5.

Since the problem is one of high speed flight, it is generally desirable to have very high natural frequencies and flutter speeds well above the normal operating speeds. Any change of stiffness or mass distribution will alter the modes and frequencies and thus allow a change in the flutter speeds. If the aircraft is not properly maintained and excessive play and flexibility exist, flutter could occur at flight speeds below the limit airspeed.

Compressibility problems may define the limit airspeed for an airplane in terms of Mach num­ber. The supersonic airplane may experience a great decay of stability at some high Mach number or encounter critical structural or engine inlet temperatures due to aerodynamic heating. The transonic airplane at an excess і ve speed may encounter a variety of stability, con­trol, or buffet problems associated with tran­sonic flight. Since the equivalent airspeed for a given Mach number decreases with altitude, the magnitude of compressibility effects at high altitude may be negligible for the tran­sonic airplane. In this sense, the airplane may not be able to fly at high enough dynamic pressures within a certain range of Mach num­bers to create any significant stability or control problem.

The transonic airplane which is buffet lim­ited requires due consideration of the effect of load factor on the onset of buffet. Since critical Mach number decreases with lift coef­ficient, the limit Mach number will decrease with load factor. If the airplane is subject to prolonged or repeated buffet for which it was not designed, structural fatigue will be the certain result.

The limit airspeed for each type aircraft is set sufficiently high that full intended appli­cation of the aircraft should be possible. Each of the factors mentioned about the effect of excess airspeed should provide due respect for the limit airspeed.

The operating flight strength limitations of afl airplane are presented in the form of a V-n or V-g diagram. This chart usually is included in the aircraft flight handbook in the section dealing with operating limitations. A typical V-n diagram is shown in figure 5-3- The V-n diagram presented in figure 5-3 is intended to present the most important general features of such a diagram and does not neces­sarily represent the characteristics of any par­ticular airplane. Each airplane type has its own particular V-n diagram with specific V’s and »’s.

The flight operating strength of an airplane is presented on a graph whose horizontal scale is airspeed (V) and vertical scale is load factor (У). The presentation of the airplane strength is contingent on four factors being known:

(1) the aircraft gross weight, (2) the configura­tion of the aircraft (clean, external stores, flaps and landing gear position, etc.), (3) symmetry of loading (since a rolling pullout at high speed can reduce the structural limits to approxi­mately two-thirds of the symmetrical load limits) and (4) the applicable altitude. A change in any one of these four factors can cause important changes in operating limits.

For the airplane shown, the positive limit load factor is 7-5 and the positive ultimate load factor is 11.25 (7.5ХІ.5)- For negative lift flight conditions the negative limit load factor is 3.0 and the negative ultimate load factor is 4.5 (3-0x15). The limit airspeed is stated as 575 knots while the wing level stall speed is apparently 100 knots.

Figure 5.4 provides supplementary informa­tion to illustrate the significance of the V-n diagram of figure 5-3- The lines of maximum lift capability are the first points of importance on the V-n diagram. The subject aircraft is capable of developing no more than one posi­tive “g" at 100 knots, the wing level stall speed of the airplane. Since the maximum load factor varies with the square of the airspeed,

STRUCTURAL

the maximum positive lift capability of this airplane is 4 “g” at 200 knots, 9 g at 300 knots, 16 g at 400 knots, etc. Any load factor above this line is unavailable aerodynamically, i. e., the subject airplane cannot fly above the line of maximum lift capability. Essentially the same situation exists for negative lift flight with the exception that the speed necessary to produce a given negative load factor is higher than that to produce the same positive load factor. Gen­erally, the negative CLmax is less than the posi­tive CLmax and the airplane may lack sufficient control power to maneuver in this direction.

If the subject airplane is flown at a positive load factor greater than the positive limit load factor of 7-5, structural damage will be possi­ble. When the airplane is operated in this region, objectionable permanent deformation of the primary structure may take place and a high rate of fatigue damage is incurred. Opera­tion above the limit load factor must be avoided in normal operation. If conditions of extreme emergency require load factors above the limit to prevent an immediate disaster, the airplane should be capable of withstanding the ultimate load factor without failure. The same situation exists in negative lift flight with the exception that the limit and ultimate load factors are of smaller magnitude and the negative limit load factor may not be the same value at all airspeeds. At speeds above the maximum level flight airspeed the negative limit load factor may be of smaller magnitude.

The limit airspeed (or redime speed) is a de­sign reference point for the airplane—the sub­ject airplane is limited to 575 knots. If flight is attempted beyond the limit airspeed struc- turahdamage or structural failure may result from a variety of phenomena. The airplane in flight above the limit airspeed may encounter: 00 critical gust (0 destructive flutter (0 aileron reversal 00 wing or surface divergence (0 critical compressibility effects such as stability and control problems, damaging buffet, etc

The occurrence of any one of these items could cause structural damage or failure of the pri­mary structure. A reasonable accounting of these items is required during the design of an airplane to prevent such occurrences in the re­quired operating regions. The limit airspeed of an airplane may be any value between termi­nal dive speed and 1.2 times the maximum level flight speed depending on the aircraft type and mission requirement. Whatever the resulting limit airspeed happens to be, it deserves due respect.

Thus, the airplane in flight is limited to a regime of airspeeds and g’s which do not exceed the limit (or redline) speed, do not exceed the limit load factor, and cannot exceed the maximum lift capability. The airplane must be operated within this “envelope” to prevent structural damage and ensure that the anticipated service life of the airplane is obtained. The pilot must appreciate the V-n diagram as describing the allowable combination of airspeeds and load factors for safe operation. Any maneuver, gust, or gust plus maneuver outside the structural envelope can cause structural damage and effectively shorten the service life of the airplane.

There are two points of great importance on the V-n diagram of figure 5-4. Point В is the intersection of the negative limit load factor and line of maximum negative lift capability. Any airspeed greater than point В provides a negative lift capability sufficient to damage the airplane; any airspeed less than point В does not provide negative lift capability sufficient to damage the airplane from excessive flight loads. Point A is the intersection of the positive limit load factor and the line of maximum positive lift capa­bility. The airspeed at this point is the minimum airspeed at which the limit load can be developed aerodynamically. Any air­speed greater than point A provides a positive lift capability sufficient to damage the air­plane; any airspeed less than point A does not provide positive lift capability sufficient to

cause damage from excessive flight loads. The usual term given to the speed at point A is the “maneuver speed,” since consideration of subsonic aerodvnamics would predict mini­mum usable turn radius to occur at this con­dition. The maneuver speed is a valuable reference point since an airplane operating below this point cannot produce a damaging positive flight load. Any combination of maneuver and gust cannot create damage due to excess airload when the airplane is below the maneuver speed.

The maneuver speed can be computed from the following equation:

VP= Va-Jn limit

where

VP= maneuver speed

Vs=stall speed

n limit = limit load factor

Of course, the stall speed and limit load factor must be appropriate for the airplane gross weight. One notable fact is that this speed, once properly computed, remains a constant value if no significant change takes place in the spanwise weight distribution. The ma­neuver speed of the subject aircraft of figure 5.4. would be

Vp=10QtJT5

= 274 knots

FLIGHT LOADS-MANEUVERS AND GUSTS

The loads imposed on an aircraft in flight are the result of maneuvers and gusts. The maneuver loads may predominate in the design of fighter airplanes while gust loads may predominate in the design of the large multiengine aircraft. The maneuver loads an airplane may encounter depend in great part on the mission type of the airplane. However, the maximum maneuvering capability is of interest because of the relationship with strength limits.

The flight load factor is defined as the pro­portion between airplane lift and weight, where

n = L/W n — load factor L=lift, lbs.

W= weight, lbs.

MANEUVERING LOAD FACTORS. The maximum lift attainable at any airspeed occurs when the airplane is at CLjnax. With the use of the basic lift equation, this maximum lift is expressed as:

Lmax = CLmaxpV2S

Since maximum lift must be equal to the weight at the stall speed,

W-Ct’J’MS

If the effects of compressibility and viscosity on Ctmax are neglected for simplification, the maximum load factor attainable is determined by the following

ГThus, if the airplane is flying at twice the stall speed and the angle of attack is increased to obtain maximum lift, a maximum load factor of four will result. At three times the stall speed, nine "g s” would result; four times the stall speed, sixteen g’s result; five times the stall speed, twenty-five g s result; etc. Therefore, any airplane which has high speed performance may have the capability of high maneuvering load factors. The airplane which is capable of flight speeds that are

many times the stall speed will require due consideration of the operating strength limits.

The structural design of the aircraft must consider the possibility of negative load factors from maneuvers. Since the pilot cannot com­fortably tolerate large prolonged negative "g", the aircraft need not be designed for negative load factors as great as the positive load factors.

The effect of airplane gross weight during maneuvers must be appreciated because of the particular relation to flight operating strength limitations. During flight, the pilot appre­ciates the degree of a maneuver from the inertia forces produced by various load factors; the airplane structure senses the degree of a maneuver principally by the airloads involved. Thus, the pilot recognizes load factor while the structure recognizes only load. To better understand this relationship, consider an ex­ample airplane whose basic configuration gross weight is 20,000 lbs. At this basic configura­tion assume a limit load factor for symmetrical flight of 5 6 and an ultimate load factor of 8.4. If the airplane is operated at any other con­figuration, the load factor limits will be al­tered. The following data illustrate this fact by tabulating the load factors required to produce identical airloads at various gross weights.

As illustrated, at high gross weights above the basic configuration weight, the limit and ulti­mate load factors may be seriously reduced. For the airplane shown, a 5-g maneuver im­mediately after a high gross weight takeoff could be very near the “disaster regime,” especially if turbulence is associated with the maneuver. In the same sense, this airplane at very low operating weights below that of the basic configuration would experience great­ly increased limit and ultimate load factors.

Operation in this region of high load factors at low gross weight may create the impression that the airplane has great excess strength capability. This effect must be understood and intelligently appreciated since it is not uncom­mon to have a modern airplane configuration with more than 50 percent of its gross weight as fuel.

GUST LOAD FACTORS. Gusts are asso­ciated with the vertical and horizontal velocity gradients in the atmosphere. A horizontal gust produces a change in dynamic pressure on the airplane but causes relatively small and unimportant changes in flight load factor. The more important gusts are the vertical gusts which cause changes in angle of attack. This process is illustrated in figure 5-2. The vec­torial addition of the gust velocity to the air­plane velocity causes the change in angle of attack and change in lift. The change in angle of attack at some flight condition causes a change in the flight load factor. The incre­ment change in load factor due to the vertical gust can be determined from the following equation:

д«=0.П5 —VL – V’QLU)

(w/s) K

where

An=change in load factor due to gust те = lift curve slope, unit of per degree of a

<r = altitude density ratio W/S= wing loading, psf Ve = equivalent airspeed, knots KU= equivalent sharp edged gust velocity ft. per sec.

As an example, consider the case of an air­plane with a lift curve slope те = 0.08 and wing loading, (WjS’) = 6Q psf. If this airplane were flying at sea level at 350 knots and encountered an effective gust of 30 ft. per sec., the gust would produce a load factor increment of 1.61. This increment would be added to the flight load factor of the airplane prior to the gust,

e. g., if in level flight before encountering the gust, a final load factor of 1.0+1.61 = 2.61 would result. As a general requirement all airplanes must be capable of withstanding an approximate effective +30 ft. per sec. gust when at maximum level flight speed for normal rated power. Such a gust intensity has rela­tively low frequency of occurrence in ordinary flying operations.

The equation for gust load increment pro­vides a basis for appreciating many of the variables of flight. The gust load increment varies directly with the equivalent sharp edged gust velocity, KU, since this factor effects the change in angle of attack. The highest reasonable gust velocity that may be anticipated is an actual vertical velocity, U, of 50 ft. per sec. This value is tempered by the fact that the airplane does not effectively encounter the full effect because of the response
of the airplane and the gradient of the gust. A gust factor, К (usually on the order of 0.6), reduces the actual gust to the equivalent sharp edged gust velocity, KU.

The properties of the airplane exert a power­ful influence on the gust increment. The lift curve slope, m, relates the sensitivity of the airplane to changes in angle of attack. An aircraft with a straight, high aspect ratio wing would have a high lift curve slope and would be quite sensitive to gusts. On the other hand, the low aspect ratio, swept wing airplane has a low lift curve slope and is com­paratively less sensitive to turbulence. The apparent effect of wing loading, WjS, is at times misleading and is best understood by considering a particular airplane encountering a fixed gust condition at various gross weights. If the airplane encounters the gust at lower than ordinary gross weight, the accelerations

due to the gust condition are higher. This is explained by the fact that essentially the same lift change acts on the lighter mass. The high accelerations and inertia forces magnify the impression of the magnitude of turbulence. If this same airplane encounters the gust condition at higher than ordinary gross weight, the accelerations due to the gust condition are lower, i. e., the same lift change acts on the greater mass. Since the pilot primarily senses the degree of turbulence by the resulting accelerations and inertia forces, this effect can produce a very misleading impression.

The effect of airspeed and altitude on the gust load factor is important from the stand­point of flying operations. The effect of alti­tude is related by the term д/o-, which would related that an airplane flying at a given EAS at 40,000 ft. ((7 = 0.25) would experience a gust load factor increment only one-half as great as at sea level. This effect results be­cause the true airspeed is twice as great and only one-half the change in angle of attack occurs for a given gust velocity. The effect of airspeed is illustrated by the linear variation of gust increment with equivalent airspeed. Such a variation emphasises the effect of gusts at high flight speeds and the probability of structural damage at excessive speeds in turbu­lence.

The operation of any aircraft is subject to specific operating strength limitations. A single large overstress may cause structural failure or damage severe enough to require costly overhaul. Less severe overstress re­peated for sufficient time will cause fatigue cracking and require replacement of parts to prevent subsequent failure. A combat airplane need not be operated in a manner like the “little old lady from Pasadena" driving to church on Sunday but each aircraft type has strength capability only specific to the mission require­ment. Operating limitations must be given due regard.