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.

Flight Control System

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.

Introduction of Missile Control System

INTRODUCTION

The missile flight control system is one element of the overall homing loop. Figure 1 is a simplified block diagram of the missile homing loop configured for the terminal phase of flight when the missile is approaching intercept with the target. The missile and target motion relative to inertial space can be combined mathematically to obtain the relative motion between the missile and the target. The terminal sensor, typically an RF or IR seeker, measures the angle between an inertial reference

and the missile-to-target line-of-sight (LOS) vector, which is called the LOS angle. The state estimator, e.g., a Kalman filter, uses LOS angle measurements to estimate LOS angle rate and perhaps other quantities such as target acceleration. The state estimates feed a guidance law that develops the flight control commands required to intercept the target. The flight control system forces the missile to track the guidance commands, resulting in the achieved missile motion. The achieved missile motion alters the relative geometry, which then is sensed and used to determine the next set of flight control commands, and so on. This loop continues to operate until the missile intercepts the target. 

In the parlance of feedback control, the homing loop is a feedback control system that regulates the LOS angle rate to zero. As such, the overall stability and performance of this control system are determined by the dynamics of each element in the loop. Consequently, the flight control system cannot be designed in a vacuum. Instead, it must be designed in concert with the other elements to meet overall homing-loop performance requirements in the presence of target manoeuvre and other disturbances in the system, e.g., terminal sensor noise (not shown in Fig. 1), which can negatively impact missile performance.

Figure 1. The flight control system is one element in the missile homing loop.

 The inertial missile motion controlled by the flight control system 

combines with the target motion to form the relative geometry 

between the missile and target. The terminal sensor measures

 the missile-to-target LOS angle. The state estimator forms 

an estimate of the LOS angle rate, which in turn is input 

to the guidance law. The output of the guidance law is the 

steering command, typically a translational acceleration.The flight 

control system uses the missile control effectors, such as 

aerodynamic tail surfaces, to force the missile to track steering 

commands to achieve a target intercept.