PITCH-UP

The term of "pitch-up” generally applies to the static longitudinal instability encountered by certain configurations at high angle of attack. The condition of pitch-up is illustrated by the graph of CM versus CL in figure 4.33. Positive static longitudinal stability is evident at low values of CL by the negative slope of the curve. At higher values of Ct the curve changes to a positive slope and large positive pitching moments are developed. This sort of in­stability implies that an increase in angle of attack produces nose up moments which tend to bring about further increases in angle of attack hence the term "pitch-up” is applied.

There are several items which may con­tribute to a pitch-up tendency. Sweepback of the wing planform can contribute unstable moments when separation or stall occurs at the tips first. The combination of sweepback and taper alters the lift distribution to produce high local lift coefficients and low energy boundary layer near the tip. Thus, the tip stall is an inherent tendency of such a plan – form. In addition, if high local lift coefficients exist near the tip, the tendency will be to incur the shock induced separation first in these areas. Generally, the wing will contribute to pitch-up only when there is large sweepback.

Of course, the wing is not the only item con­tributing to the longitudinal stability of the airplane. Another item important as a source of pitch-up is the down wash at the horizontal tail. The contribution of the tail to stability depends on the change in tail lift when the air­plane is given a change in angle of attack. Since the downwash at the tail reduces the change in angle of attack at the tail, any in­crease in downwash at the tail is destabilizing.

For certain low aspect ratio airplane configura­tions, an increase in airplane angle of attack may physically locate the horizontal tail in

Mach number. As a corollary of this increase in stability is a decrease in controllability and an increase in trim drag.

The static directional stability of an air­plane decreases with Mach number in super­sonic flight. The influence of the fuselage and the decrease in vertical tail lift curve slope bring about this condition.

The dynamic stability of the airplane generally deteriorates with Mach number in supersonic flight. Since a large part of the damping depends on the tail surfaces, the decrease in lift curve slope with Mach number will account in part for the decrease in damp­ing. Of course, all principal motions of the aircraft must have satisfactory damping and if the damping is not available aerodynami­cally it must be provided synthetically to obtain satisfactory flying qualities. For many high speed configurations the pitch and yaw dampers, flight stabilization systems, etc., are basic necessities rather than luxuries.

Generally, flight at high Mach number will take place at high altitude hence the effect of high altitude must be separated for study. All of the basic aerodynamic damping is due to moments created by pitching, rolling, or yawing motion of the aircraft. These moments are derived from the changes in angles of attack on the tail surfaces with angular rotation (see fig. 4.15). The very high true airspeeds common to high altitude flight reduce the angle of attack changes and reduce the aerodynamic damping. In fact, the aero­dynamic damping is proportional to •фГ, similar to the proportion of true airspeed to equivalent airspeed. Thus, at the altitude of

40,0 ft., the aerodynamic damping would be reduced to one-half the sea level value and at the altitude of 100,000 ft. the aerodynamic damping would be reduced to one-tenth the sea level value.

High dynamic pressures (high q) can be common to flight at high Mach number and adverse aeroelastic effects may be encountered. If the aircraft surfaces encounter significant deflection when subject to load, the tendency may be to lower the contribution to static stability and reduce the damping contribution. Thus, the problem of adequate stability of the various airplane motions is aggravated.