What is Yaws?

Yaw is a short-time rotational movement of the aircraft around the vertical axis of the airframe coordinate system. The vertical axis passes through the center of gravity of the aircraft in the plane of symmetry of the aircraft and is perpendicular to the longitudinal axis. The main parameters of yaw motion are yaw angle, yaw angular velocity, and yaw angular acceleration. The yaw motion is realized by stepping on the pedal to yaw the rudder to generate a yaw moment about the center of gravity of the aircraft. ^[1]

The main control surface for yaw is the vertical tail with rudder. When determining the size of the vertical tail, the following points need to be considered.

(1) The size of the vertical tail must be able to adapt to the range of gravity center movement required in the entire flight envelope area.

(2) In the event of an engine accident, especially a wing engine, the vertical tail must be able to generate sufficient lateral force to balance this unstable moment.

(3) In landing mode, vertical tail size is often limited by crosswind requirements.

For the above engine stopping conditions, for an aircraft with a wing engine, the take-off state is usually a critical indicator for determining the size of the vertical tail. ^[3]

Similar to the roll and pitch methods, the acceleration feedback of the yaw attitude is constructed using the speed gyroscope, and the stability of the system has been significantly improved. Therefore, it is possible to fully control the yaw angle by a simple control method . Therefore, a P control system is designed outside the angular velocity feedback system, as shown in the overview diagram. In the overview diagram, K _p is the gain of P and _ref refers to the yaw angle. Given the gyro gain, when K _p = 1, the P gain can be determined through simulation using a closed-loop loop, which can be obtained by integrating the yaw attitude model and the feedback loop using the speed gyro system. ^[4]

Similar to the roll case, the yaw maneuver evaluation is divided into the following stages, as shown in Figure 3-7.

When the aircraft is in a side-slip condition or encounters a horizontal gust, the lift distribution on the flat tail is asymmetric. This is due to the tendency of the tail and the fuselage to tilt towards the flat tail on one side. This is especially important when the flat tail is mounted on the vertical tail, because the rolling torque of the flat tail always increases the bending moment of the vertical tail. This situation is usually the critical torsional load at the rear of the fuselage. Slipstreams or jets can also produce similar asymmetric loads. ^[5]

The lateral component of the speed of the aircraft during sideslip will affect the aerodynamic characteristics of the whole aircraft, not just the aerodynamic characteristics of the tail. The wing-body zero-lift pitching moment will also change, which will cause the flat tail load to change during the aircraft's trim flight. The British military code recommends that when estimating the flat tail trim load value in a side-skid maneuver, the zero-lift pitch moment coefficient is increased by -0.0015 / side-slip angle compared to a straight flight. ^[5]