The flight control system is one element of the overall homing loop. Figure 2 shows the basic elements of the flight control system, which itself is another feedback control loop within the overall homing loop depicted in Fig. 1. An inertial measurement unit (IMU) measures the missile translational acceleration and angular velocity. The outputs of the IMU are combined with the guidance commands in the autopilot to compute the commanded control input, such as a desired tail-surface deflection or thrust-vector angle. An actuator, usually an electromechanical system, forces the physical control input to follow the commanded control input. The airframe dynamics respond to the control input. The basic objective of the flight control system is to force the achieved missile dynamics to track the guidance commands in a well-controlled manner. The figures of merit (FOMs) used to assess how well the flight control system works are discussed in Flight Control System Design Objectives.
GUIDANCE INPUT:
The inputs to the flight control system are outputs from the guidance law that need to be followed to ultimately effect a target intercept. The specific form of the flight control system inputs (acceleration commands, attitude commands, etc.) depends on the specific application . In general, the flight control system must be designed based on the expected characteristics of the commands, which are determined by the other elements of the homing loop and overall system requirements. Characteristics of concern can be static, dynamic, or both. An example of a static characteristic
is the maximum input that the flight control system is expected to be able to track. For instance, a typical rule of thumb for intercepting a target that has constant acceleration perpendicular to the LOS is for the missile to have a 3:1 acceleration advantage over the target. If the missile system is expected to intercept a 10-g accelerating threat, then the flight control system should be able to force the missile to maintain a 30-g acceleration. An example of a dynamic characteristic is the expected frequency content of the command. For instance, rapid changes in the command are expected as the missile approaches intercept against a maneuvering threat, but the input commands may change more slowly during 11the midcourse phase of flight where the objective is to keep the missile on an approximate collision path or to minimize energy loss. Other dynamic characteristics of concern include the guidance command update rate and the amount of terminal sensor noise flowing into the flight control system and causing unnecessary control actuator activity.
AIRFRAME DYNAMICS:
Recall that the objective of the flight control system is to force the missile dynamics to track the input command. The dynamics of the airframe are governed by fundamental equations of motion, with their specific characteristics determined by the missile aerodynamic response, propulsion, and mass properties. Assuming that missile motion is restricted to the vertical plane (typical for early concept development), the equations of motion that govern the missile dynamics can be developed
in straightforward fashion.
ACTUATOR:
The missile actuator converts the desired control command developed by the autopilot into physical motion, such as rotation of a tail fin, that will effect the desired missile motion. Actuators for endoatmospheric missiles typically need to be high-bandwidth devices (significantly higher than the desired bandwidth of the flight control loop itself) that can overcome significant loads. Most actuators are electromechanical, with hydraulic actuators being an option in certain applications. Although the actuator often is modeled as a linear system for preliminary design and development, it is actually a nonlinear device, and care must be taken by the flight control designer not to exceed the hardware capabilities. Two critical FOMs for the actuator for many endoatmospheric missiles are its rate and position limits. The position limit is an effective limit on the moment that the control input can impart on the airframe, which in turn limits the maximum AOA (mentioned in Figure 5) and acceleration. The rate limit essentially limits how fast the actuator can cause the missile to rotate, which effectively limits how fast the flight control system can respond to changes in the guidance command. The performance of a flight control system that commands the actuator to exceed its limits can be degraded, particularly if the missile is flying at a condition where it is statically unstable.
INERTIAL MEASUREMENT UNIT (IMU):
The IMU measures the missile dynamics for feedback to the autopilot. In most flight control applications, the IMU is composed of accelerometers and gyroscopes to measure three components of the missile translational acceleration and three components of missile angular velocity. Like the actuator, the IMU needs to be a high-bandwidth device relative to the desired bandwidth of the flight control loop. In some applications, other quantities also need to be measured, such as the pitch angle for an attitude control system. In this case, other sensors can be used (e.g., an inertially stabilized platform), or IMU outputs can feed strapdown navigation equations that are implemented in a digital computer to determine the missile attitude, which then is sent to the autopilot as a feedback measurement. The flight control system must be designed such that the missile dynamics do not exceed the dynamic range of the IMU. If the IMU saturates, the missile will lose its inertial reference, and the flight control feedback is corrupted. The former may be crucial, depending on the specific missile application and the phase of flight. The latter may be more problematic if the dynamic range is exceeded for too long, particularly if the missile is statically unstable.
AUTOPILOT:
The autopilot is a set of equations that takes as inputs the guidance commands and the feedback measurements from the IMU and computes the control command as the output. As mentioned previously, the autopilot must be designed so that the control command does not cause oversaturation of the actuator or the IMU. Because the autopilot usually is a set of differential equations, computing its output involves integrating signals with respect to time. Most modern autopilots are implemented in discrete time on digital computers, although analog autopilots are still used. The following section describes several types of autopilots that apply in different flight control applications.
GUIDANCE INPUT:
The inputs to the flight control system are outputs from the guidance law that need to be followed to ultimately effect a target intercept. The specific form of the flight control system inputs (acceleration commands, attitude commands, etc.) depends on the specific application . In general, the flight control system must be designed based on the expected characteristics of the commands, which are determined by the other elements of the homing loop and overall system requirements. Characteristics of concern can be static, dynamic, or both. An example of a static characteristic
is the maximum input that the flight control system is expected to be able to track. For instance, a typical rule of thumb for intercepting a target that has constant acceleration perpendicular to the LOS is for the missile to have a 3:1 acceleration advantage over the target. If the missile system is expected to intercept a 10-g accelerating threat, then the flight control system should be able to force the missile to maintain a 30-g acceleration. An example of a dynamic characteristic is the expected frequency content of the command. For instance, rapid changes in the command are expected as the missile approaches intercept against a maneuvering threat, but the input commands may change more slowly during 11the midcourse phase of flight where the objective is to keep the missile on an approximate collision path or to minimize energy loss. Other dynamic characteristics of concern include the guidance command update rate and the amount of terminal sensor noise flowing into the flight control system and causing unnecessary control actuator activity.
AIRFRAME DYNAMICS:
Recall that the objective of the flight control system is to force the missile dynamics to track the input command. The dynamics of the airframe are governed by fundamental equations of motion, with their specific characteristics determined by the missile aerodynamic response, propulsion, and mass properties. Assuming that missile motion is restricted to the vertical plane (typical for early concept development), the equations of motion that govern the missile dynamics can be developed
in straightforward fashion.
ACTUATOR:
The missile actuator converts the desired control command developed by the autopilot into physical motion, such as rotation of a tail fin, that will effect the desired missile motion. Actuators for endoatmospheric missiles typically need to be high-bandwidth devices (significantly higher than the desired bandwidth of the flight control loop itself) that can overcome significant loads. Most actuators are electromechanical, with hydraulic actuators being an option in certain applications. Although the actuator often is modeled as a linear system for preliminary design and development, it is actually a nonlinear device, and care must be taken by the flight control designer not to exceed the hardware capabilities. Two critical FOMs for the actuator for many endoatmospheric missiles are its rate and position limits. The position limit is an effective limit on the moment that the control input can impart on the airframe, which in turn limits the maximum AOA (mentioned in Figure 5) and acceleration. The rate limit essentially limits how fast the actuator can cause the missile to rotate, which effectively limits how fast the flight control system can respond to changes in the guidance command. The performance of a flight control system that commands the actuator to exceed its limits can be degraded, particularly if the missile is flying at a condition where it is statically unstable.
INERTIAL MEASUREMENT UNIT (IMU):
The IMU measures the missile dynamics for feedback to the autopilot. In most flight control applications, the IMU is composed of accelerometers and gyroscopes to measure three components of the missile translational acceleration and three components of missile angular velocity. Like the actuator, the IMU needs to be a high-bandwidth device relative to the desired bandwidth of the flight control loop. In some applications, other quantities also need to be measured, such as the pitch angle for an attitude control system. In this case, other sensors can be used (e.g., an inertially stabilized platform), or IMU outputs can feed strapdown navigation equations that are implemented in a digital computer to determine the missile attitude, which then is sent to the autopilot as a feedback measurement. The flight control system must be designed such that the missile dynamics do not exceed the dynamic range of the IMU. If the IMU saturates, the missile will lose its inertial reference, and the flight control feedback is corrupted. The former may be crucial, depending on the specific missile application and the phase of flight. The latter may be more problematic if the dynamic range is exceeded for too long, particularly if the missile is statically unstable.
AUTOPILOT:
The autopilot is a set of equations that takes as inputs the guidance commands and the feedback measurements from the IMU and computes the control command as the output. As mentioned previously, the autopilot must be designed so that the control command does not cause oversaturation of the actuator or the IMU. Because the autopilot usually is a set of differential equations, computing its output involves integrating signals with respect to time. Most modern autopilots are implemented in discrete time on digital computers, although analog autopilots are still used. The following section describes several types of autopilots that apply in different flight control applications.
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