LONGITUDINAL STABILITY AND CONTROL

STATIC LONGITUDINAL STABILITY

GENERAL CONSIDERATIONS. An air­craft will exhibit positive static longitudinal stability if it tends to return to the trim angle of attack when displaced by a gust or control movement. The aircraft which is unstable will continue to pitch in the disturbed direction until the displacement is resisted by opposing control forces. If the aircraft is neutrally stable, it tends to remain at any displacement to which it is disturbed. It is most necessary to provide an airplane with positive static longitudinal stability. The stable airplane is safe and easy to fly since the airplane seeks and tends to maintain a trimmed condition of flight. It also follows that control deflec­tions and control “feel’’ are logical in direction
and magnitude. Neutral static longitudinal stability usually defines the lower limit of. airplane stability since it’ is the boundary between stability and instability. The air­plane with neutral static stability may be excessively responsive to controls and the aircraft has no tendency to return to trim fol­lowing a disturbance. The airplane with negative static longitudinal stability is in­herently divergent from any intended trim condition. If it is at all possible to fly the aircraft, the aircraft cannot be trimmed and illogical control forces and deflections are re­quired to provide equilibrium with a change of attitude and airspeed.

Since static longitudinal stability depends upon the relationship of angle of attack and pitching moments, it is necessary to study the pitching moment contribution of each com­ponent of the aircraft. In a manner similar to all other aerodynamic forces, the pitching

moment about the lateral axis is studied in the coefficient form.

M = pitching moment about the c. g., ft.- lbs., positive if in a nose-up directioti q~ dynamic pressure, psf i’=wing area, sq. ft.

МЛС = mean aerodynamic chord, ft.

CM—pitching moment coefficient

The pitching moment coefficients contributed by all the various components of the aircraft are summed up and plotted versus lift coeffi­cient. Study of this plot of CM versus CL will relate the static longitudinal stability of the airplane.

Graph A of figure 4.5 illustrates the variation of pitching moment coefficient, CM, with lift coefficient, CL, for an airplane with positive static longitudinal stability. Evidence of static stability is shown by the tendency to re­turn to equilibrium—or “trim"—upon dis­placement. The airplane described by graph A is in trim or equilibrium when CM = 0 and, if the airplane is disturbed to some different CL, the pitching moment change tends to return the aircraft to the. point of trim. If the airplane were disturbed to some higher CL (point У), a negative or nose-down pitching moment is de­veloped which tends to decrease angle of attack back to the trim point. If the airplane were disturbed to some lower CL (point X), a posi­tive) or nose-up pitching moment is developed which tends to increase the angle of attack back to the trim point. Thus, positive static longitudinal stability is indicated by a negative slope of CM versus CL, i. e., positive stability is evidenced by a decrease in Cm with an increase in Ch.

The degree of static longitudinal stability is indicated by the slope of the curve of pitching moment coefficient with lift coefficient. Graph

В of figure 4.5 provides comparison of the stable and unstable conditions. Positive sta­bility is indicated by the curve with negative slope. Neutral static stability would be the result if the curve had zero slope. If neutral stability exists, the airplane could be dis­turbed to «оте higher or lower lift coefficient without change in pitching moment coefficient. Such a condition would indicate that the air­plane would have no tendency to return to some original equilibrium and would not hold trim. An airplane which demonstrates a posi­tive slope of the CM versus CL curve would be unstable. If the unstable airplane were subject to any disturbance from equilibrium at the trim point, the changes in pitching moment would only magnify the disturbance. When the unstable airplane is disturbed to some higher Cl, a positive change in CM occurs which would illustrate a tendency for continued, greater displacement. When the unstable air­plane is disturbed to some lower CL, a negative change in Cm takes place which tends to create continued displacement.

Ordinarily, the static longitudinal stability of a conventional airplane configuration does not vary with lift coefficient. In other words, the slope of CM versus CL does not change with Cb – However, if the airplane has sweepback, large contribution of power effects to stability, or significant changes in downwash at the horizontal tail, noticeable changes in static stability can occur at high lift coefficients. This condition is illustrated by graph C of figure 4.5. The curve of CM versus Cl of this illustration shows a good stable slope at low values of Сь. Increasing CL effects a slight decrease in the negative slope hence a decrease in stability occurs. With continued increase in CL, the slope becomes zero and neutral stability exists. Eventually, the slope be­comes positive and the airplane becomes un­stable or “pitch-up" results. Thus, at any lift coefficient, the static stability of the air­plane is depicted by the slope of the curve of CM versus CL.

CONTRIBUTION OF THE COMPONENT SURFACES. The net pitching moment about the lateral axis is due to the contribution of each of the component surfaces acting in their appropriate flow fields. By study of the con­tribution of each component the effect of each component on the static stability may be ap­preciated. It is necessary to recall that the pitching moment coefficient is defined as:

Thus, any pitching moment coefficient—re­gardless of source—has the common denomi­nator of dynamic pressure, q, wing area, S, and wing mean aerodynamic chord, MAC. This common denominator is applied to the pitch­ing moments contributed by the fuselage and nacelles, horizontal tail, and power effects as well as pitching moments contributed by the wing.

WING. The contribution of the wing to stability depends primarily on the location of the aerodynamic center with respect to the airplane center of gravity. Generally, the aerodynamic center—or a. c.—is defined as the point on the wing mean aerodynamic chord where the wing pitching moment coefficient does not vary with lift coefficient. All changes in lift coefficient effectively take place at the wing aerodynamic center. Thus, if the wing experiences some change in lift coefficient, the pitching moment created will be a direct function of the relative location of the a. c. and

c-g-

Since stability is evidenced by the develop­ment of restoring moments, the c. g. must be forward of the a. c. for the wing to contribute to positive static longitudinal stability. As shown in figure 4.6, a change in lift aft of the c. g. produces a stable restoring moment de­pendent upon the lever arm between the a. c. and c. g. In this case, the wing contribution would be stable and the curve of CM versus CL for the wing alone would have a negative slope. If the c. g. were located at the a. c., CM would not vary with CL since all changes in lift would take place at the c. g. In this case, the wing contribution to stability would be neutral. When the c. g. is located behind the a. c. the wing contribution is unstable and the curve of Cu versus CL for the wing alone would have a positive slope.

Since the wing is the predominating aero­dynamic surface of an airplane, any change in the wing contribution may produce a sig­nificant change in the airplane stability. This fact would be most apparent in the case of the flying wing or tailless airplane where the wing contribution determines the airplane stability. In order for the wing to achieve stability, the c. g. must be ahead of the a. c. Also, the wing must have a positive pitching moment about the aerodynamic center to achieve trim at positive lift coefficients. The first chart of figure 4.7 illustrates that the wing which is stable will trim at a negative lift coefficient if the CuAC is negative. If the stable wing has a positive С**лс it will then trim at a useful posi­tive CL. The only means available to achieve trim at a positive CL with a wing which has a negative CUac is an unstable c. g. position aft of the a. c. As a result, the tailless aircraft cannot utilize high lift devices which incur any significant changes in С**лс-

While the trim lift coefficient may be altered by a change in c. g. position, the resulting change in stability is undesirable and is unsat­isfactory as a primary means of control. The variation of trim CL by deflection of control surfaces is usually more effective and is less inviting of disaster. The early attempts at manned flight led to this conclusion.

When the aircraft is operating in subsonic flight, the a. c. of the wing remains fixed at the 25 percent chord station. When the aircraft is flown in supersonic flight, the a. c. of the wing will approach the 50 percent chord sta­tion. Such a large variation in the location of the a. c. can produce large changes in the wing contribution and greatly alter the air­plane longitudinal stability. The second chart

Figure 4.6. Wing Contribution

Figure 4.7. Effect of Cmac* C. G. Position and Mach Niimber

of figure 4.7 illustrates the change of wing contribution possible between subsonic and supersonic flight. The large increase in static stability in supersonic flight can incur high trim drag or require great control effectiveness to prevent reduction in maneuverability.

FUSELAGE AND NACELLES. In most cases, the contribution of the fuselage and nacelles is destabilizing. A symmetrical body of revolution in the flow field of a perfect fluid develops an unstable pitching moment when given an angle of attack. In fact, an increase in angle of attack produces an increase in the unstable pitching moment without the devel­opment of lift. Figure 4.8 illustrates the pres­sure distribution which creates this unstable moment on the body of revolution. In the actual case of real subsonic flow essentially the same effect is produced. An increase in angle of attack causes an increase in the unstable pitching moment but a negligible increase in lift.

An additional factor for consideration is the influence of the induced flow field of the wing. As illustrated in figure 4.8, the upwash ahead of the wing increases the destabilizing influence from the portions of the fuselage and nacelles ahead of the wing. The downwash behind the wing reduces the destabilizing influence from the portions of the fuselage and nacelles aft of the wing. Hence, the location of the fuselage and nacelles relative to the wing is important in determining the contribution to stability.

The body of revolution in supersonic flow can develop lift of a magnitude which cannot be neglected. When the body of revolution in supersonic flow is given an angle of attack, a pressure distribution typical of figure 4.8 is the result. Since the center of pressure is well forward, the body contributes a destabilizing influence. As is usual with supersonic con­figurations, the fuselage and nacelles may be quite large in comparison with the wing area and the contribution to stability may be large. Interaction between the wing and fuselage and nacelles deserves consideration in several in­stances. Body upwash and variation of local Mach number can influence the wing lift while lift carryover and downwash can effect the fu­selage and nacelles forces and moments.

HORIZONTAL TAIL. The horizontal tail usually provides the greatest stabilizing influ­ence of all the components of the airplane. To appreciate the contribution of the horizontal tail to stability, inspect figure 4-9- If the air­plane is given a change in angle of attack, a change in tail lift will occur at the aerody­namic center of the tail. An increase in lift at the horizontal tail produces a negative moment about the airplane e g. and tends to return the airplane to the trim condition. While the contribution of the horizontal tail to stability is large, the magnitude of the Contribution is dependent upon the change in tail lift and the lever arm of the surface. It is obvious that the horizontal tail will produce a stabilizing effect only when the surface is aft of the c. g. For this reason it would be inap­propriate to refer to the forward surface of a canard (tail-first) configuration as a horizontal “stabilizer.” In a logical sense, the horizontal “stabilizer” must be aft of the c. g. and— generally speaking—the farther aft, the greater the contribution to stability.

Many factors influence the change in tail lift which occurs with a change in airplane angle of attack. The area of the horizontal tail has the obvious effect that a large surface would generate a large change in lift. In a similar manner, the change in tail lift would depend on the slope of the lift curve for the horizontal tail. Thus, aspect ratio, taper, sweepback, and Mach number would deter­mine the sensitivity of the surface to changes in angle of attack. It should be appreciated that the flow at the horizontal tail is not of the same flow direction or dynamic pressure as the free stream. Due to the wing wake, fuse­lage boundary layer, and power effects, the q at the horizontal tail may be greatly different from the q of the free stream. In most ІП-

BODY OF REVOLUTION IN PERFECT FLUID

INDUCED FLOW FIELD FROM WING

stances, the q at the tail is usually less and this reduces the efficiency of the tail.

When the airplane is given a change in angle of attack, the horizontal tail does not expe­rience the same change in angle of attack as the wing. Because of the increase in down – wash behind the wing, the horizontal tail will experience a smaller change in angle of attack, e. g., if a 10° change in wing angle of attack causes a 4° increase in downwash at the hori­zontal tail, the horizontal tail experiences only a 6° change in angle of attack. In this manner, the downwash at the horizontal tail reduces the contribution to stability. Any factor which alters the rate of change of down – wash at the horizontal tail will directly affect the tail contribution and airplane stability.

Power effects cah alter the downwash at the horizontal tail and affect the tail contribution. Also, the downwash at the tail is affected by the lift distribution on the wing and the flow condition on the fuselage. The low aspect ratio airplane requires large angles of attack to achieve high lift coefficients and this posi­tions the fuselage at high angles of attack. The change in the wing downwash can be accompanied by crossflow separation vortices on the fuselage. It is possible that the net effect obviates or destabilizes the contribu­tion of the horizontal tail and produces air­plane instability.

POWER-OFF STABILITY. When the in­trinsic stability of a configuration is of interest, power effects are neglected and the stability is considered by a buildup of the contributing components. Figure 4.10 illustrates a typical buildup of the components of a conventional airplane configuration. If the c. g. is arbi­trarily set at 30 percent MAC, the contribu­tion of the wing alone is destabilizing as indi­cated by the positive slope of CM versus CL. The combination of the wing and fuselage increases the instability. The contribution of the tail alone is highly stabilizing from the large negative slope of the curve. The contribution of the tail must be sufficiently stabilizing so that the complete configuration will exhibit positive static stability at the anticipated c. g. locations. In addition, the tail and wing incidence must be set to provide a trim lift coefficient near the design condition.

When the configuration of the airplane is fixed, a variation of c. g. position can cause large changes in the static stability. In the conventional airplane configuration, the large changes in stability with c. g. variation are primarily due to the large changes in the wing contribution. If the incidence of all surfaces remains fixed, the effect of c. g. position on static longitudinal stability is typified by the second chart of figure 4.10. As the c. g. is gradually moved aft, the airplane static sta­bility’ decreases, then becomes neutral then unstable. The c. g. position which produces zero slope and neutral static stability is re­ferred to as the “neutral point.” The neutral point may be imagined as the effective aerody­namic center of the entire airplane configura – ration, i. e., with the c. g. at this position, all changes in net lift effectively occur at this point and no change in pitching moment results. The neutral point defines the most aft c. g. position without static instability.

POWER EFFECTS. The effects of power may cause significant changes in trim lift coefficient and static longitudinal stability. Since the contribution to stability is evaluated by the change in moment coefficients, power effects will be most significant when the airplane operates at high power and low airspeeds such as the power approach or waveoff condition.

The effects of power are considered in two main categories. First, there are the direct effects resulting from the forces created by the propulsion unit. Next, there are the indirect effects of the slipstream and other associated flow which alter the forces and moments of the aerodynamic surfaces. The direct effects of power are illustrated in figure 4.11. The ver­tical location of the thrust line defines one of the direct contributions to stability. If the

thrust line is below the c. g., thrust produces a positive or noseup moment and the effect is de­stabilizing. On the other hand, if the thrust line is located above the c. g., a negative moment is created and the effect is stabilizing.

A propeller or inlet duct located ahead of the c. g. contributes a destabilizing effect. As shown in figure 4.11, a rotating propeller in­clined to the windstream causes a deflection of the airflow. The momentum change of the
slipstream creates a normal force at the plane of the propeller similar to a wing creating lift by deflecting an airstream. As this normal force will increase with an increase in airplane angle of attack, the effect will be destabilizing when the propeller is ahead of the c. g. The magnitude of the unstable contribution de­pends on the distance from the c. g. to the propeller and is largest at high power and low dynamic pressure. The normal force created

EFFECT OF VERTICAL LOCATION OF THRUST LINE

at the inlet of a jet engine contributes a similar destabilizing effect when the inlet is ahead of the c. g. As with the propeller, the magni­tude of the stability contribution is largest at high thrust and low flight speed.

The indirect effects of power are of greatest concern in the propeller powered airplane rather than the jet powered airplane. As shown in figure 4.12, the propeller powered airplane creates slipstream velocities on the various surfaces which are different from the flow field typical of power-off flight. Since the various wing, nacelle, and fuselage surfaces are partly or wholly immersed in this slip­stream, the contribution of these components to stability can be quite different from the power-off flight condition. Ordinarily, the change of fuselage and nacelle contribution with power is relatively small. The added lift on the portion of the wing immersed in the slipstream requires that the airplane oper­ate at a lower angle of attack to produce the same effective lift coefficient. Generally, this reduction in angle of attack to effect the same CL reduces the tail contribution to stability. However, the increase in dynamic pressure at the tail tends to increase the effectiveness of the tail and may be a stabilizing effect. The magnitude of this contribution due to the slipstream velocity on the tail will depend on the c. g. position and trim lift coefficient.

The deflection of the slipstream by the nor­mal force at the propeller tends to increase the down wash at the horizontal tail and reduce the contribution to stability. Essentially the same destabilizing effect is produced by the flow induced at the exhaust of the jet power – plant. Ordinarily, the induced flow at the horizontal tail of a jet airplane is slight and is destabilizing when the jet passes underneath the horizontal tail. The magnitude of the indirect power effects on stability tends to be greatest at high Ct, high power, and low flight speeds.

The combined direct and indirect power effects contribute to a general reduction of static stability at high power, high and low q. It is generally true that any airplane will experience the lowest level of static longi­tudinal stability under these conditions. Be­cause of the greater magnitude of both direct and indirect power effects, the propeller pow­ered airplane usually experiences a greater effect than the jet powered airplane.

An additional effect on stability can be from the extension of high lift devices. The high lift devices tend to increase downwash at the tail and reduce the dynamic pressure at the tail, both of which are destabilizing. However, the high lift devices may prevent an unstable contribution of the wing at high CL. While the effect of high lift devices depends on the airplane configuration, the usual effect is de­stabilizing. Hence, the airplane may experi­ence the most critical forward neutral point during the power approach or waveoff. Dur­ing these conditions of flight the static stability is usually the weakest and particular attention must be given to precise control of the air­plane. The power-on neutral point may set the most aft limit of c. g. position.

CONTROL FORCE STABILITY. The static longitudinal stability of an airplane is defined by the tendency to return to equilibrium upon displacement. In other words, the stable air­plane will resist displacement from the trim or equilibrium. The control forces of the air­plane should reflect the stability of the air­plane and provide suitable reference for precise control of the airplane.

The effect of elevator deflection on pitching moments is illustrated by the first graph of figure 4.13. If the elevators of the airplane are fixed at zero deflection, the resulting line of Cm versus CL for 0° depicts the static stability and trim lift coefficient. If the elevators are fixed at a deflection of 10° up, the airplane static stability is unchanged but the trim lift coefficient is increased. A change in elevator or stabilizer position does not alter the tail contribution to stability but the change in pitching moment will alter the lift coefficient

at which equilibrium will occur. As the ele­vator is fixed in various positions, equilibrium (or trim) will occur at various lift coefficients and the trim CL can be correlated with elevator deflection as in the second graph of figure 4.13.

When the c. g. position of the airplane is fixed, each elevator position corresponds to a particular trim lift coefficient. As the c. g. is moved aft the slope of this line decreases and the decrease in stability is evident by a given control displacement causing a greater change in trim lift coefficient. This is evidence that decreasing stability causes increased controlla­bility and, of course, increasing stability de­creases controllability. If the c. g. is moved aft until the line of trim CL versus elevator de­flection has zero slope, neutral static stability is obtained and the “stick-fixed” neutral point is determined.

Since each value of lift coefficient corresponds to a particular value of dynamic pressure re­quired to support an airplane in level flight, trim airspeed can be correlated with elevator deflection as in the third graph of figure 4.13. If the c. g. location is ahead of the stick-fixed neutral point and control position is directly related to surface deflection, the airplane will give evidence of stick position stability. In other words, the airplane will require the stick to be moved aft to increase the angle of attack and trim at a lower airspeed and to be moved forward to decrease the angle of attack and trim at a higher airspeed. To be sure, it is desirable to have an airplane demon­strate this feature. If the airplane were to have stick position instability, the airplane would require the stick to be moved aft to trim at a higher airspeed or to be moved forward to trim at a lower airspeed.

There may be slight differences in the static longitudinal stability if the elevators are allowed to float free. If the elevators are allowed to float free as in “hands-off"’ flight, the elevators may have a tendency to “float” or streamline when the horizontal tail is given a change in angle of attack. If the horizontal tail is subject to an increase in angle of attack and the elevators tend to float up, the change in lift on the tail is less than if the elevators remain fixed and the tail contribution to stability is reduced. Thus, the “stick-free” stability of an airplane is usually less than the stick-fixed stability. A typical reduction of stability by free elevators is shown in figure 4.14(A) where the airplane, stick-free demon­strates a reduction of the slope of CM versus CL. While aerodynamic balance may be provided to reduce control forces, proper balance of the surfaces will reduce floating and prevent great differences between stick-fixed and stick-free stability. The greatest floating tendency oc­curs when the surface is at a high angle of attack hence the greatest difference between stick-fixed and stick-free stability occurs when the airplane is at high angle of attack.

If the controls are fully powered and actu­ated by an irreversible mechanism, the sur­faces are not free to float and there is no differ­ence between the stick-fixed and stick-free static stability.

The control forces in a conventional air-

і

plane are made up of two components. First, the basic stick-free stability of the airplane contributes an increment, of force which is independent of airspeed. Next, there, is an increment of force dependent on the trim tab sttting which varies with-the dynamic pres­sure or the square of equivalent airspeed. Figure 4.14(B) indicates the variation of stick force with airspeed and illustrates the effect of tab setting on stick force. In order to trim the airplane at point (1) a certain amount of up elevator is required and zero stick force is obtained with the use of the tab. To trim the airplane for higher speeds corre­sponding to points (2) and (3) less and less nose-up tab is required. Note that when the airplane is properly trimmed, a push force is required to increase airspeed and a pull force is required to decrease airspeed. In this man­ner, the airplane would indicate positive stick force stability with a stable “feel” for air-

speed. If the airplane were given a large nose down tab setting the pull force would in­crease with airspeed. This fact points out the possibility of “feel” as not being a true indi­cation of airplane static stability.

If the c. g. of the airplane were varied while maintaining trim at a constant airspeed, the effect of c. g. position on stick force stability could be appreciated. As illustrated in figure 4.14(C), moving the c. g. aft decreases the slope of the line of stick force through the trim speed. Thus, decreasing stick force stability is evident in that smaller stick forces are necessary to displace the airplane from the trim speed. When the stick force gradient (or slope) becomes zero, the c. g. is at the stick-free neutral point and neutral stability exists. If the c. g. is aft of the stick-free neutral point, stick force instability will exist, c. g. the airplane will require a push force at a lower speed or a pull force at a higher speed. It should be noted that the stick force gradient is low at low airspeeds and when the airplane is at low speeds, high power, and a c. g. position near the aft limit, the “feel” for airspeed will be weak.

Control system friction can create very un­desirable effects on control forces. Figure 4.14(D) illustrates that the control force versus airspeed is a band rather than a line. A wide friction force band can completely mask the stick force stability when the stick force stability is low. Modern flight control systems require precise maintenance to mini­mize the friction force band and preserve proper feel to the airplane.

MANEUVERING STABILITY. When an airplane is subject to a normal acceleration, the flight path is curved and the airplane is subject to a pitching velocity. Because of the pitching velocity in maneuvering flight, the longitudinal stability of the airplane is slightly greater than in steady flight condi­tions. When an airplane is subject to a pitch­ing velocity at a given lift coefficient, the air­plane develops a pitching moment resisting the pitch motion which adds to the restoring moment from the basic static stability. The principal source of this additional pitching moment is illustrated in figure 4.15.

During a pull-up the airplane is subject to an angular rotation about the lateral axis and the horizontal tail will experience a component of wind due to the pitching velocity. The vector addition of this component velocity to the flight velocity provides a change in angle of attack for the tail and the change in lift on the tail creates a pitching moment resisting the pitching motion. Since the pitching mo­ment opposes the pitching motion but is due to the pitching motion, the effect is a damping in pitch. Of course, the other components of the airplane may develop resisting moments and contribute to pitch damping but the horizontal tail is usually the largest contri­bution. The added pitching moment from pitch damping will effect a higher stability in maneuvers than is apparent in steady flight. From this consideration, the neutral point for maneuvering flight will be aft of the neutral point for unaccelerated flight and in most cases will not be a critical item. If the airplane demonstrates static stability in unaccelerated flight, it will most surely demonstrate stability in maneuvering flight.

The most direct appreciation of the ma­neuvering stability of an airplane is obtained from a plot of stick force versus load factor such as shown in figure 4.15. The airplane with positive maneuvering stability should demonstrate a steady increase in stick force with increase in load factor or “G”. The maneuvering stick force gradient—or stick force per G—must be positive but should be of the proper magnitude. The stick force gradient must not be excessively high or the airplane will be difficult and tiring to maneuver. Also, the stick force gradient must not be too low or the airplane may be overstressed in­advertently when light control forces exist. A maneuvering stick force gradient of 3 to 8 lbs. per G is satisfactory for most fighter and

Figure 4.15. Maneuvering Stability

attack airplanes. A large patrol or transport type airplane would ordinarily show a much higher maneuvering stick force gradient be­cause of the lower limit load factor.

When the airplane has high static stability, the maneuvering stability will be high and a high stick force gradient will result. A possibility exists that the forward c. g. limit could be set to prevent an excessively high maneuvering stick force gradient. As the c. g. is moved aft, the stick force gradient de­creases with decreasing maneuvering stability and the lower limit of stick force gradient may be reached.

The pitch damping of the airplane is obvi­ously related to air density. At high altitudes, the high true airspeed reduces the change in tail angle of attack for a given pitching velocity and reduces the pitch damping. Thus, a de­crease in maneuvering stick force stability can be expected with increased altitude.

TAILORING CONTROL FORCES. The control forces should reflect the stability of the airplane but, at the same time, should be of a tolerable magnitude. The design of the surfaces and control system may employ an infinite variety of techniques to provide satis­factory control forces.

Aerodynamic balance must be thought of in two different senses. First, the control surface must be balanced to reduce hinge moments due to changes in angle of attack. This is necessary to reduce the floating tendency of the surface which reduces the stick-free stability. Next, aerodynamic balance can reduce the hinge moments due to deflection of the control sur­face. Generally, it is difficult to obtain a high degree of deflection balance without incurring a large overbalance of the surface for changes in angle of attack.

Some of the types of aerodynamic balance are illustrated in figure 4.16. The-simple horn type balance employs a concentrated balance area located ahead of the hinge line. The balance area may extend completely to the leading edge (unshielded) or part’way to the leading edge (shielded). Aerodynamic balance can be achieved by the provision of a hinge line aft of the control surface leading edge. The resulting overhang of surface area ahead of the hinge line will provide a degree of balance depending on the amount of overhang. Another variation of aerodynamic balance is afl internal balance surface ahead of the hinge line which is contained within the surface. A flexible seal is usually incorporated to in­crease the effectiveness of the balance area. Even the bevelling of the trailing edge of the control surface is effective also as a balancing technique. The choice of the type of aerody­namic balance will depend on many factors such as required degree of balance, simplicity, drag, etc.

Many devices can be added to a control system to modify or tailor the stick force stability to desired levels. If a spring is added to the control system as shown in figure 4.16, it will tend to center the stick and provide a force increment depending on stick displace­ment. When the control system has a fixed gearing between stick position and surface deflection, the centering spring will provide a contribution to stick force stability according to stick position. The contribution to stick force stability will be largest at low flight speeds where relatively large control deflec­tions are required. The contribution will be smallest at high airspeed because of the smaller control deflections required. Thus, the stick centering bungee will increase the airspeed and maneuvering stick force stability but the contribution decreases at high airspeeds. A variation of this device would be a spring stiffness which would be controlled to vary with dynamic pressure, q – In this case, the contribution of the spring to stick force stability would not diminish with – speed.

A “downspring” added to a control system is a means of increasing airspeed stick force stability without a change in airplane static

TYPES OF AERODYNAMIC BALANCE

OVERHANG OR LEADING EDGE BALANCE BY OFFSET HINGE

C

d

EFFECT OF STICK CENTERING SPRING

EFFECT OF DOWNSPRING

stability. As shown in figure 4.17, a down­spring consists of a long preloaded spring at­tached to the control system which tends to rotate the elevators down. The effect of the downspring is to contribute an increment of pull force independent of control deflection or airspeed. When rhe downspring is added to the control system of an airplane and the air­plane is retrimmed for the original speed, the airspeed stick force gradient is increased and there is a stronger feel for airspeed. The down­spring would provide an “ersatz” improve­ment to an airplane deficient in airspeed stick force stability. Since the force increment from the downspring is unaffected by stick position or normal acceleration, the maneuvering stick force stability would be unchanged.

The bobweight is an effective device for im­proving stick force stability. As shown in figure 4.17, the bobweight consists of an eccen­tric mass attached to the control system which—in unaccelerated {light—contributes an increment of pull force identical to the downspring. In fact, a bobweight added to the control system of an airplane produces an effect identical to the downspring. The bob – weight will increase the airspeed stick force gradient and increase the feel for airspeed.

A bobweight will have an effect on the maneuvering stick force gradient since the bob – weight mass is subjected to the same accelera­tion as the airplane. Thus, the bobweight will provide an increment of stick force in direct proportion to the maneuvering acceleration of the airplane. Because of the linear contribu­tion of the bobweight, the bobweight can be applied to increase the maneuvering stick force stability if the basic airplane has too low a value or develops a decreasing gradient at high lift coefficients.

The example of the bobweight is useful to point out the effect of the control system dis­tributed masses. All carrier aircraft must have the control system mass balanced to prevent undesirable control forces from the longi­tudinal accelerations during catapult launching.

Various control surface tab devices can be utilized to modify control forces. Since the de­flection of a tab is so powerful in creating hinge moments on a control surface, the possible application of tab devices is almost without limit. The basic trim tab arrangement is shown in figure 4.18 where a variable linkage connects the tab and the control surface. Ex­tension or contraction of this linkage will de­flect the tab relative to the control surface and create a certain change in hinge moment coef­ficient. The use of the trim tab will allow the pilot to reduce the hinge moment to zero and trim the control forces to zero for a given flight condition. Of course, the trim tab should have adequate effectiveness so that control forces can be trimmed out throughout the flight speed range.

The lagging tab arrangement shown in figure

4.18 employs a linkage between the fixed sur­face and the tab surface. The geometry is such that upward deflection of the control surface displaces the tab down relative to the control surface. Such relative displacement of the tab will aid in deflection of the control surface and thus reduce the hinge moments due to deflection. An obvious advantage of this device is the reduction of deflection hinge moments without a change in aerodynamic balance.

The leading tab arrangement shown in figure

4.18 also employs a linkage between the fixed surface and the tab surface. However, the geometry of the linkage is such that upward deflection of the control surface displaces the tab up relative to the control surface. This relationship serves to increase the control sur­face hinge moments due to deflection of the surface.

The servo tab shown in figure 4.18 utilizes a horn which has no direct connection to the control surface and is free to pivot about the hinge axis. However, a linkage connects this free horn to the tab surface. Thus, the control system simply deflects the tab and the resulting hinge moments deflect the control surface.

SERVO TAB

,HORN FREE TO PIVOT ON HINGE AXIS

^SPRING

-HORN FIXED TO SURFACE

Since the only control forces are those of the tab, this device makes possible the deflection of large surfaces with relatively small control

forces.

A variation of the basic servo tab layout is the spring tab arrangement of figure 4.18. When the control horn is connected to the control surface by springs, the function of the tab is to provide a given portion of the required control forces. The spring tab arrangement can then function as a boost to reduce control forces. The servo tab and spring tab are usually applied to large or high speed subsonic airplanes to provide tolerable stick forces.

The spring loaded tab of figure 4.18 consists of a free tab preloaded with a spring which furnishes a constant moment about the tab hinge line. When the airplane is at zero air­speed, the tab is rotated up to the limit of deflection. As airspeed is increased, the aero­dynamic hinge moment on the tab will finally equal the spring torque and the tab will begin to streamline. The effect of this arrangement is to provide a constant hinge moment to the control system and contribute a constant push force requirement at speeds above the preload speed. Thus, the spring loaded tab can im­prove the stick force gradient in a manner similar to the downspring. Generally, the spring loaded tab may be more desirable because of greater effectiveness and the lack of undesirable control forces during ground operation.

The various tab devices have almost un­limited possibilities for tailoring control forces. However, these devices must receive proper care and maintenance in order to function properly. In addition, much care must be taken to ensure that no slop or play exists in the joints and fittings, otherwise destructive flutter may occur.