Stability Augmentation System (SAS) - Engineering Seminar

Introduction

Stability augmentation system (SAS) were generally the first feedback control system design intended to improve dynamic stability characteristic. There were also referred to as damper, stabilizers, and stability augmenters. Aircraft such as the F-104 had SAS. This system generally fed back an aircraft motion parameter, such as pitch rate, to provide a control deflection that opposed the motion and increase damping characteristic. The SAS had to be integrated with the primary mechanical control system of the aircraft consisting of the stick, pushrods, cables, and bellcranks leading to the control surface or the hydraulic actuator that activated the control surface. The control authority (percentage of full surface deflection avaible)

 of SAS was generally limited to about 10 %. Figure 9.1 present a simplified SAS.One problem with SAS was the fact that feedback loop provided a command that oppose pilot control input. As a result, the aircraft become less responsive for a given stick input. This was typical addressed with the addition of a washout filter in the feedback loop that attenuated the feedback signal for constant value of the aircraft motion parameter.

Another concern was the limited authority of the SAS actuator that was necessitated by safety-of-flight requirement. SAS sensor and computers were normally nonredundant or duel redundant and thus did not approach the system reliability of the mechanical flight control system. Despite these concerns, SAS was effective in improving aircraft flying qualities.    

Understanding of Stability Augmentation System (SAS)

Stability augmentation system is system that alters of the inherent behaviour of a system. This system is used on aircraft. This is usually achieved by a system which controls one or more flight-control surfaces (or engines) automatically without inputs from the pilot. The inherent stability and response behaviour of many modern airplanes tends toward low damping or even instability. The physical reasons have to do with the configuration of the airplane and the combination of flight speed and altitude at which the airplane is operated. Several modern fighters and even some transports are intentionally designed with no or little inherent stability. There are a number of reasons for such a design condition. In the case of fighters, excellent manoeuvre ability in combat is essential. By making a fighter intentionally inherently unstable, it is easy to design the control system so that load factors in pull-ups or in turns can be built up rapidly. In the case of transports, the motivation to design for little or no inherent stability is to lower the size of the tail and thereby achieve a reduction in drag and weight

            The control exercised by the stability augmentation system contrasts with that exercised by the pilot. The pilot may be connected with the flight-control surface via a direct mechanical link. Alternatively, in many modern airplanes the pilot cockpit control movement is sensed by a position transducer. The output of the position transducer in turn is sent, via a computer-amplifier combination, to a hydraulic actuator, referred to as a servo, which drives the flight-control surface.

Command signals which come from the pilot or from the stability augmentation system are sent by wire (fly-by-wire) or by optical conduit (fly-by-light) to the electromagnetic valve. A valve distributes high-pressure hydraulic fluid either to the left or to the right of the piston so that the piston is forced to move. The piston in turn moves the flight-control surface.With the introduction of fast in-flight digital computers, it has become possible to equip airplanes with so-called full flight envelop protection systems. Such systems are designed to refuse any pilot input which might get the airplane into a flight condition from which recovery is no longer possible. Such systems can easily be arranged to prevent a pilot from rolling a commercial airplane too much or to prevent the pilot from stalling the airplane. Such systems can also be arranged so that loads acting on the wing or tail do not approach dangerously high levels. In that case the system is referred to as a load-alleviation system.

A functional block diagram of a typical flight control (FCS) is shown in fig.0. It is assume that the primary flying controls are mechanical such that pilot commands drive the control surface via control actuators which augment that available power to level sufficient to overcome the aerodynamic loads on the surface. The electronic flight control system (EFCS) comprises two feedback loop both of which derive their control signal from motion sensors appropriate to the requirement of the control laws. The output from the inner and outer loop controller is electronically summed and the resultant signals control the aircraft via a small servo actuator. Typically, the servo actuator is an electro-hydraulic device which convert low power electrical signal to mechanical signal at a power level compatible with those originated at the pilot to which it is mechanical summed. Although only a single control axes is indicated in figure 1, it is important to appreciate that the FCS will, in general, include closed loop controllers operating on the roll, pitch and yaw control axes of the aircraft simultaneously and may even extend to include closed loop engine control as well. Thus multi-variable feedback involve many separate control loop is implied, which is typical of many modern FCS.

The inner loop provide stability augmentation and is usually regarded as essential for continued proper operation for aircraft. The inner loop control system alone comprise the stability augmentation system (SAS), it is usually the first part of the FCS to be design and together with the airframe comprises the augmented aircraft.

  The outer loops provide the autopilot which, as its name suggest, enable the pilot to fly various manoeuvre under automatic control. Although necessary for operational reason, an autopilot is not essential for the provision of a save well behave aircraft. The autopilot control mode are designed to function with the augmented aircraft and may  be selective engaged as required to automated the piloting task. Their use is intend to release the pilot from monotony of flying steady condition which may be at, or beyond, limit of human capability. Autopilot control modes vary from the very simple, for example height hold, to the very complex, for example automatic landing.

            Since, typically, for most aircraft the control law gains required to effects good stability, control and handling vary with operating condition, it is necessary to make provision for their continuous adjustment. The variations often arise as a result of variations in the aerodynamics properties of airframes over the flight envelope. For example, at low speed the aerodynamics effectiveness of the control surfaces is generally less than at high speed. This means that higher control gains are required at low speeds and vice versa. It is, therefore, common practice to vary, or schedule, gains as a function of flight control. Commonly use flight condition variables are dynamics pressure, Mach number, altitude and so on, information which is grouped under the description of air data. Generally, air data information would be available to all control laws in a FCS as indicated in Fig. 1

            A control panel is provided in the cockpit to enable the pilot to control and monitor the operations of the FCS. SAS controls are usually minimal and enable the pilot to monitor the system for correct, and hence safe, operation. In some cases he may also be provided with means for selectively isolating parts of the SAS. On the other hands, the autopilot control panels is rather more substantial. Controls are provided to enable the pilot set up, engage and disengage the various autopilot mode functions. The control panel also enables him to monitor progress during the automated manoeuvres selected.

            In piloted phases of flight the autopilot would normally be disengaged and, as indicated in Fig. 0 the pilot would derive his perception of flying and handling qualities from the motion cues provided by the augmented aircraft. Thus the inner loop control system provided the means by which all aspects of stability,  control and handling may be tailored in order to improve the characteristics of the basic aircraft.

The Control Law

the control law is a mathematical expression which describes the function implemented by an augmentation or autopilot controller. For example, a very simple and very commonly used control law describing an inner loop control system for augmenting yaw damping is

            ζ(s)=K_ζδ_ζ(s) – K_r(s/1+sT)r(s)      

Equation above simply states that the control signal applied to the rudder ζ(s) comprises the sum of the pilot command δ_ζ(s) and yaw rate feedback r(s). The gain K_ζ is the mechanical gearing between rudder pedals and rudder and the gain K_r is the all important feedback gain chosen by design to optimise the damping in yaw. The second term in equation above  is negative since negative feedback is required to increase stability in yaw. The second term also, typically, includes a washout, or high-pass, filter with a time constant of around 1 or 2 s. the filter is included to block yaw rate feedback in steady turning flight in order to prevent the feedback loop opposing the pilot command once the rudder pedals are returned to centre after manoeuvre initiation. However, the filter is effectively transparent during transient motion thereby enabling the full effect of the feedback loop quickly damp out the yaw oscillation.

For Dynamic and Static Stability

The main goal Stability augmentation systems (SAS) areto make the aircraft more stable. There are SASs are divide into two parts that are dynamic stability (whether the eigen motions don’t diverge) and the static stability (whether the equilibrium position itself is stable).

Dampers – Acquiring dynamic stability.

An airplane has several eigenmotions. When the properties of these eigenmotions don’t comply with the requirements, we need an SAS. The SAS is mostly used to damp the eigenmotions. Therefore, we will now examine how various eigenmotions are damped. For this part,SAS consist of  :

1.      The yaw damper: modelling important systems

2.      The yaw damper: determining the transfer function

3.      The pitch damper

4.      The phugoid damper

1. The yaw damper: modelling important systems

When an aircraft has a low speed at a high altitude, the Dutch roll properties of the aircraft deteriorate. To prevent this, a yaw damper is used. An overview of this system. The yaw damper gets its input (feedback) from the yaw rate gyro. It then sends a signal to the rudder servo. The rudder is then moved in such a way that the Dutch roll is damped much more quickly than usual. As a designer, the only parts that can be influence is the yaw damper.

 However, it is not necessary to know how other systems work as well. For this reason, those model systems are made. Usually the model of the aircraft is assume to be known. (Or use the one that is derived in the Flight Dynamics course.) So, it will only examine the other systems.

The time constant Tservo depends on the type of actuator. For slow electric actuators,    Tservo  0.25. However, for fast hydraulic actuators, Tservo  0.05 to 0.1. This time constant (or equivalently, the servo break frequency servo ) can be very important. If it turns out to be different than expected, the results can also be very different. So, it is often worth while to investigate what happens if Tservo varies a bit.

2.The yaw damper: determining the transfer function

As we know that the yaw damper has to reduce the yaw rate. But it shouldn’t always try to keep the yaw rate at zero. In this case, the pilot will have a hard time to change the heading of the aircraft. Thus, a reference yaw rate r is also supplied to the system. This yaw rate can be calculated from the desired heading rate by using.

This will cause the yaw damper to fight less when a yaw rate is continuously present. In other words, the system ‘adjusts’ itself to a new desired yaw rate. The time constant  is quite important. For too high values, the pilot will still have to fight the yaw damper. But for too low values, the yaw damper itself doesn’t work, because the washout circuit simply adjusts too quickly. A good compromise is oftenat  = 4s.

Finally, at the yaw damper transfer function, it  have proportional, integral and derivative action. If the rise time should be reduced, proportional action is choosen. If the steady state error needs to be reduced, add an integral action. And if the transient, response needs to be reduced (e.g. to reduce overshoot) we apply a derivative action. In this way, the right values of Kp, KI and KD can be chosen.

Sometimes, the optimal values of the gains Kp, KI and KD differ per flight phase. In this case gain scheduling can be applied. The gains then depend on certain relevant parameters, like the velocity V and the altitude h. In this way, every flight phase will have the right gains.

3.The pitch damper

When an aircraft flies at a low speed and a high altitude, the short period eigenmotion has a low damping. To compensate for this, a pitch damper is used. The pitch damper is in many ways similar to the yaw damper. Also the set-up is similar. Only this time, the elevators and a pitch rate gyro are used, instead of the rudder and a yaw rate gyro.

Alternatively, a washout circuit can again be used. This washout circuit again has the function given in equation above. Also, a value of   = 4 is again a good compromise. Just like a yaw damper, also the pitch damper has proportional, integral and derivative actions.

4. The phugoid damper

To adjust the properties of the phugoid, phugoid damper can be used. It is very similar to the previous two dampers have been seen. However, this damper uses the measured velocity U as input. Its output is sent to the elevator. The speed sensor and the elevator servo are modelled as

Note that for the servo now a break frequency of ωbr = 20 Hz is assumed.

A reference velocity U is also needed by the system. This reference velocity is simply set by the pilot/autopilot. Alternatively, a washout circuit can be used. This washout circuit is the same as those of the yaw and pitch damper. And, just like the previous two dampers, again proportional, integral and derivative actions can be used.When using a phugoid damper, one should also keep in mind the short period motion properties. Improvingthe phugoid often means that the short period properties become worse.

Acquiring static stability -Feedback

Before an aircraft can be dynamically stable, it should first be statically stable. In other words, shouldhave C < 0 and C > 0. Normal aircraft already have this. But very manoeuvrable aircraft, like fighter aircraft, do not. (less stability generally means more manoeuvrability). For this part,  SAS consist of  :

1.      Angle of attack feedback

2.      Load factor feedback

3.      Sideslip feedback

1. Angle of attack feedback

To make an aircraft statically stable, feedback is applied. The most important part is the kind of feedback that is used. First, angle of attack feedback for longitudinal control. In other words, the angle of attack   is used as a feedback parameter. First, model the angle of attack sensor and the (canard) servo actuator. This is often done using,

So, now a break frequency  = 40 is used for the servo.

For angle of attack feedback, usually only a proportional gain K_ is used. By using the models of the sensor and actuator (and of course also the aircraft), a root locus plot can be made. With this root locus plot, a nice value of the gain K can be chosen. This gain is then used to determine the necessary canard  deflection canard. This is done using

canard = K

However, a check does need to be performed on whether the canard deflections can be achieved. If gust loads can cause a change in angle of attack of  = 1and the maximum canard deflection is 25, then K  should certainly not be bigger than 25, or even be close to it for that matter.

2. Load factor feedback

There is a downside with angle of attack feedback. It is often hard to measure
 accurately. So instead, load factor feedback can be applied. Now the value of n is used as feedback. As models for the sensor and actuator,  again use

Need a model for the aircraft. Normally, it assume that such a model is known. However, the transfer function between the load factor n and the canard deflection c is usually not part of the aircraft model. So, it simply derive it. For that, we first can use

Then divide the equation by c. If it also use   = , then it find that

The transfer functions from c to both  and  usually are part of the aircraft model. So it can be assume that they are known. The transfer function between n and c is thus now also known. All that is left for us to do is choose an appropriate gain Kn. And of course, again it needs to be checked whether this gain Kn doesn’t result into too big canard deflections. The load factor sensor also has a downside. It is often hard to distinguish important accelerations (like the ones caused by turbulence) from unimportant accelerations (like vibrations due to, for example, a firing gun). Good filters need to be used to make sure a useful signal is obtained.

3. Sideslip feedback

Previously it already  considered longitudinal stability. For lateral stability, sideslip feedback can be used. (However, sideslip feedback is not yet applied in practice.) With sideslip feedback, the sideslip angle is used as feedback parameter for the rudder. The sensor and the rudder are usually modelled as

The transfer function between the sideslip angle  and the rudder deflection r usually follows from the airplane model. Now that the model is in place, a nice gain K  can be chosen for the system. This should then give it the right properties.

There is a small problem with sideslip feedback. It can generate a lateral phugoid mode of vibration. To compensate for this, another feedback loop is often used, where the roll rate is used as feedback for the ailerons. This then reduces the effects of the lateral phugoid motion.

Safety of Stability Augmentation System

Stability Augmentation System is one of the parts of Flight Control System (FCS).In FCS there are two loops which is inner and outer loop. The inner loop provides the SAS and is usually regarded essential for continued proper operation of the flight. The inner loop control system alone comprises the SAS, it is usually the first part of FCS to be designed and together with the airframe comprises the augmented aircraft.

In any aircraft fitted with a FCS safety is the most critical part. Since the FCS has direct access to the control surfaces considerable care must e exercised in the design of the system to ensure that under no circumstances can a maximum instantaneous uncontrolled command be applied to any control surfaces.

SAS is intended to enhance safety by reducing pilot workload. It is not a substitute for adequate pilot skill nor does it relieve the pilot of the responsibility to maintain adequate outside visual reference. If unexpected attitude deviations occur, and/or the cyclic forces and/or motions are erratic, the pilot should take manual control of the cyclic and disengage SAS immediately. With SAS engaged, pilot must always monitor the flight controls and aircraft attitude, and be prepared to immediately assume full manual control if required.