Types Of Flight Control System

The specific type of flight control system that is implemented on a particular missile depends on several factors, including the overall system mission and requirements, packaging constraints, and cost. In many applications, the type of flight control system changes with different phases of flight. For example, the system used during the boost phase for a ground- or ship-launched missile could very well differ from the system used during the intercept phase. This section provides a brief overview of different types of flight control systems and when they might be used.

1- Acceleration Control System
2- Attitude Control System
3- Flight Path Angle Control System

1- ACCELERATION CONTROL SYSTEM:

One type of flight control system common in many endoatmospheric applications is designed to track commanded acceleration perpendicular to the missile longitudinal axis. In this system, deflection of an aerodynamic control surface, such as a tail fin, is the control input, and pitch angular rate (q) and acceleration (Az) are measured by the IMU for feedback to the autopilot. The control deflection produces a small aerodynamic force on the tail fin but a large moment on the airframe because of its lever arm from the center of mass. The induced moment rotates the missile to produce the AOA, which in turn produces aerodynamic lift to accelerate the airframe. Figure 1 represents equations that can be implemented in the autopilot to develop the commanded control surface deflection angle d based on the commanded acceleration and feedback measurements of achieved acceleration Az and pitch rate q. This particular structure can be found in many missile applications, but is by no means exclusive. As indicated in Fig. 1, the error between the commanded and achieved acceleration is used as an input to the inner control loops that control missile pitch rate.The pitch rate control loops include integration with respect to time that is implemented in circuitry for an analog autopilot or with numerical difference equations in a computer in a digital autopilot. The three control gains are selected so that the closed-loop flight control system has the desired speed of response and robustness consistent with other design constraints such as actuator limits. The autopilot in Fig.1 is a reasonable starting point for a preliminary design. The final implementation would need to include other features, such as additional filters to attenuate IMU noise and missile vibrations, so that the system would actually work in flight and not just on paper.

Figure 1. This block diagram illustrates a classical approach to the design of an acceleration control autopilot. The difference between the scaled input acceleration command and the measured acceleration is multiplied by a gain to effectively form a pitch rate command. The difference between the effective pitch rate command and the measured pitch rate is multiplied by a gain and integrated with respect to time. The resulting integral is differenced
with the measured pitch rate and multiplied by a third gain to form the control effector command such as desired tail-deflection angle. The gain on the input acceleration command ensures zero steady-state error to constant acceleration command inputs. The final autopilot design would build on this basic structure with the addition of noise filters and other features such as actuator command limits. This basic structure is called the three-loop autopilot.

2- ATTITUDE CONTROL SYSTEM:

Figure 2 shows another type of autopilot that can be used to control the attitude of the missile. In this case, the control effector is the thrust-deflection angle that is actuated by either a nozzle or jet tabs. The feedback loops have a structure similar to that used in the acceleration control system of Fig. 1, except that the outer loop is pitch-angle feedback instead of acceleration. The numerical values of the gains in the control loops may differ for controlling attitude compared to controlling translational acceleration. The integration of pitch rate measured by the IMU to pitch attitude would typically be done via discrete integration in the missile navigation processing in the flight computer.

Figure 2. This attitude control system also has a three-loop structure

like the acceleration control system. The difference lies in the selection of the numerical values of the gains to reflect a different design criterion, i.e., controlling attitude instead of acceleration.

3- FLIGHT PATH ANGLE CONTROL SYSTEM:

Figure 3 shows an autopilot that can be used to track flight-path angle commands using thrust-vector control. This type of system assumes that aerodynamic forces are small and hence applies for exoatmospheric flight or for endoatmospheric flight when the missile speed is low. In this design, the feedback loops reflect the underlying physical relationships among the flight-path angle, AOA, flight-path angle rate, and pitch rate. The design explicitly uses estimates of the missile thrust and mass properties to compensate for how the missile dynamics change as propellant is expended. The commanded input into the autopilot is the desired flight-path angle. The output is the thrust-vector deflection angle. The feedback signals are the pitch rate q, flight-path angle rate g·, AOA a, and flight-path angle g. The pitch rate is measured by the IMU. The other feedback quantities are estimated in the missile navigation processing in the flight computer. This design is an example of the dynamic inversion design approach, which will be discussed in Technology Development.

Figure 3. This diagram of a flight-path control system shows the dynamic inversion design approach. The design explicitly uses the fundamental relationships among the missile kinematic and dynamic variables as well as real-time estimates of the missile thrust and mass properties to naturally compensate for the changing missile dynamics as propellant is expended.