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|  MIL-C-18244A(WEP)
3.1.1.4.3
Switching - Switching with zero command signal input from external guid-
ante systems shall not cause transients greater than + 0.05 g normal acceleration at the center of
gravity in pitch or + 1 degree in the roll attitude.
3.1.1.4.4
Noise Compatibility - The automatic flight control system shall be so design-
ed that the noise content in the external guidance signal, as specified in the applicable system speci-
fication, shall not saturate any component of the automatic flight control system, shall not impair
the response of the aircraft to the proper guidance signals, and shall not cause objectionable control
surface motion or attitude variation. If the specified noise content is too great to achieve this goal,
additional noise filtering shall be employed. Since additional noise filters impair the guidance per-
formance, an optimum compromise between performance and noise filtering shall be determined by
the procuring activity, the automatic flight control system contractor and the contractor responsible
for the guidance computer and the overall guidance performance.
3.1.1.4.5
Data Link - If the steering information is transmitted to the automatic flight
control system via a digital data link, the sampling frequency and number of bits per signal shall be
compatible with the accuracy and dynamic performance requirements of the guidance loop, and the
noise resulting from the sampling and digitalizing process shall not cause a total noise which will
be incompatible with 3.1.1.4.4. If the steering information is transmitted to the automatic flight
control system via an analog data link, the gain variation and the zero shift of the data link shall be
compatible with the performance and accuracy requirements of the guidance loop and the data link
noise shall not cause a total noise which will be incompatible with 3.1 1.4.4.
3.1.1.5
Performance Requirements - The aerodynamic and flight configurations,
external stores configuration, and aircraft performance range through which the automatic flight
control system shall be required to provide the specified performance shall be as defined in the
applicable specification. The performance requirements specified herein shall apply to all fixed.
wing aircraft, helicopters, and VTOL aircraft during forward flight at a speed greater than 30 knots.
Deviations from the performance requirements specified herein shall be allowed only as necessary,
and shall be subject to the approval of the procuring activity.
Augmentation - The augmentation system shall provide handling character-
3.1.1.5.1
istics which will satisfy, as a minimum, the requirements of Specification MIL-F-8785 for all fixed-
wing aircraft and VTOL aircraft in the cruise configuration and Specification MIL-H-8501 for heli-
copters and VTOL in the hover and transition configurations. During turn maneuvers, the augmenta-
tion system shall provide turn coordination as specified in 3.1.1.5.2.4. The control authority of the
augmentation system shall be limited as far as possible to insure that a "hard-over" signal will not
cause the aircraft to exceed its limit load factor. If this is not possible because of the demands of the
augmentation system, additional requirements shall be specified in the applicable system specification
to insure the safety of the weapons system operation.
3.1.1.5.2
Pilot Assist Function -
Attitude Hold (Pitch and Roll) - The selected pitch and roll attitudes shall be
3.1.1.5.2.1
maintained within a static accuracy of 0.5 degree with respect to the gyro reference. Upon comple-
tion of a pilot controlled maneuver, the time attitude maintained by the automatic flight control
system shall be the airplane attitude at the time the commanded forces were removed, if this attitude
is within the limits of the attitude hold mode. When using a flight controller, the airplane shall return
to a wings level attitude when the turn control is placed in the detent position.
3.1.1.5.2.1.1
Pitch Transient Response - The short period pitch response shall be smooth
and rapid. After the automatic flight control system has been manually overpowered to change the
pitch attitude by at least + 5 degrees, the aircraft shall return to the reference attitude within one
overshoot which shall not exceed 20 percent of the initial deviation. The period of overpowering shall
be short enough to hold the airspeed change to within 5 percent of the trim airspeed,
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