NAVWEPS 00-801-80  STABILITY AND CONlRO ’ L  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.  speed, If the airplane were given a large nose  down tab setting the pull force would in-  crease  with airspeed. This fact points out the  possibility of “feel”  as not being a true indi-  cation 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, e.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.  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 mo-  ment 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 contri-  bution.  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.  Control system friction can create very un-  desirable 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 mini-  mize the friction force band and preserve  proper feel to the airplane.  The most direct appreciation of the ma-  neuvering 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  in-  advertently when light control forces exist.  A maneuvering stick force gradient of 3 to 8  lbs. per G is satisfactory for most fighter and  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 condi-  tions.  When an airplane is subject to a pitch-  1 ing velocity at a given lift coefficient, the air-  plane develops a pitching moment resisting 