F-15 Flight Control
Having taken a look at the mechanical
aspects of the F-15 Flight Control System, let's turn our attention
to the electronic portion, the Control Augmentation System.
Possibly the most frequently asked questions are: "Just what is the
Control Augmentation System?" "What does it do for me?" "How does it
The Control Augmentation System (CAS) consists of two distinct functions. The first is our old friend, the Stability Augmentation System, otherwise known as Stab Aug or SAS. For those old-timers who can remember far enough back, this used to be called a Damper. The Stab-Aug, or Damper, portion of the F-15 CAS is designed to help stabilize the airframe, compensating for unwanted motion which might occur as a result of wind gusts or disturbances.
The second CAS function is its Control Stick Steering mode. This measures, compares, shapes, and smooths out pilot stick and pedal inputs allowing precise and comfortable control throughout the maneuvering envelope
Why is CAS desirable? Well, we know that the F-15 airframe is basically stable, and that the manual flight controls are designed to give Level II handling without augmentation. (Military Specification MIL-F-8785B de-fines Level II handling as "flying qualities adequate to accomplish the mission Flight Phase, but some increase in pilot workload or degradation in mission effectiveness, or both, exists.") Despite the basic stability of the Eagle, various flight conditions and varieties of store loadings could result in some pretty touchy handling situations were it not for the CAS. In addition, the CAS provides safe control of the aircraft should the basic mechanical system suffer failure or battle damage such as foreign object jams or shot-away linkage.
The bulk of this article will be concentrated upon the yaw and roll CAS (you'll see, shortly, that these two functions can't be separated). The pitch CAS will be the subject of a future article.
Sideslip Control and Damper - Precise sideslip control is provided during maneuvering by the yaw CAS. Referring to the yaw channel block diagram (Figure 1), you'll see a Rudder Pedal Position Linear Voltage Differential Transformer (LVDT). As the pilot applies force to the pedals, the me-chanical system begins to deflect the rudders. At the same time, the pedal position LVDT generates an electrical signal. As the aircraft responds to the pedal input, the yaw rate gyro and the lateral accelerometer will sense this motion.
Their blended signals are compared to
the pedal position LVDT
signal within the Roll/Yaw Computer and the resultant signal output
either adds or subtracts rudder control surface deflection as needed
to obtain the proper response. Likewise, this combination of signals
directs the servo amplifiers to deflect the rudder power cylinder
and rudder control surfaces the correct amount and direction to
dampen any unwanted yaw disturbance.
Elimination of Steady-State Sideslip - A circuit was added to reduce uncommanded and persistent sideslip (due primarily to rudder linkage friction or hysteresis) during supersonic flight and following a maneuver. The combined lateral acceleration and yaw rate sensor feedback voltages are compared with the rudder trim position and the combined output is used by the Proportional plus Integral (P + I) circuit to apply rudder surface deflection in a direction to eliminate the sideslip. With a maximum authority of ±3.75 degrees rudder, pilots can expect the ball to be pretty well neutral during flight.
There is, however, one disadvantage to the P + I circuit. If rudder surface hangup exceeds the P + I authority, yaw CAS cannot automatically trim the aircraft to zero sideslip. Thus, some manual pedal trim may be required to make up the difference, reducing sideslip to minimum. Manual trim should be applied slowly, or in small amounts with waiting periods, since the new pedal LVDT position affects the integration of the P + I circuit. If this procedure is not followed, it may appear to the pilot that he is chasing the trim.
Pilots can expect to see yaw trim changes varying in magnitude with different aircraft anytime the landing gear is extended or yaw CAS is disengaged. Both of these actions drive the P + I integrator to zero, introducing a yaw transient. So we may have solved the supersonic sideslip problem but created a problem at yaw CAS shutdown.
Maintenance personnel will have to locate and reduce system friction to a minimum. A good source of system friction can be found in the flexible cables. Here are a few points to keep in mind:
• Any kinks or rough spots are
cause for cable replacement.
• At any attach point such as bellcranks or ARI output rods, attempt to line up the cable end exactly with its attach point. In other words, reduce any apparent side load; the nature of the ribbon cable is to increase friction loads as side force is exerted on the ends.
• Make changes in direction with as large a radius as possible with only minor twists.
To summarize, problems will go away when mechanical rudder linkage is kept friction-free.
Failure Detection and Shutdown -
All three CAS channels utilize a dual-channel "Fail-Off" system. There are several things that would cause the failure detection circuitry to shut down CAS: if spin is evident (yaw rate exceeds 41.5 degrees/second); a malfunction unbalances the yaw computation circuits; an imbalance between the rudder actuators; and failure of the rudder actuator shut-off valves. Let's amplify a bit on these situations -
• If CAS is causing or aggravating the spin mode, we want CAS off. Therefore, any yaw rate in excess of 41.5 degrees/second will cause yaw CAS to shut down. Roll shuts down as a result of yaw shutting down and pitch follows if the high yaw rate continues for a period of longer than 120 milliseconds.
• Yaw CAS will shut down for malfunctions which unbalance the yaw computation circuits. This may result from differences in one of the dual yaw rate gyros or lateral accelerometer output voltages which exceed a preset level. The system will also be shut down by an electronic failure within the yaw computation circuits.
• A shutdown of yaw CAS will occur if a problem in the system causes one rudder hydraulic actuator to mistrack the other by approximately four degrees for a period of one second or longer.
• The final cause for shutdowns applies to production computers Part Number 275E514G3 which will be installed effective F-15 ship 61 and TF-15 ship 14. Circuitry has been added to these computers which senses open wiring to the rudder surface actuator shutoff valve or a failure of the actuator shutoff valve itself. In the older C2 computers, an open in shutoff valve wiring renders the yaw CAS inoperative; however, the roll CAS will not be shut down and the roll and yaw CAS telepanel lights remain out. Because of the inoperative shut-off valve, electrical signals from the CAS will not affect the rudders; the only rudder movement will come from mechanical inputs. Since the rudders will not mistrack in this mode, there will be no shutdown.
During preflights (maintenance and aircrew), ground personnel should double check to insure that yaw CAS inserts an additional 50 percent (or 15 degrees) of rudder surface deflection, making a total surface movement of 30 degrees (these figures are approximate).
Aileron Rudder Interconnect - In order to improve turn
coordination, roll rate signals are applied to the yaw channel. Since greater rudder
deflection is required for turn coordination at high angles of attack, this roll
rate signal is scheduled with angle of attack, increasing rudder
deflection as angle of attack is increased. To minimize roll and
yaw coupling tendencies, ARI is defeated at Mach numbers above
1.5 and for negative angles of attack by driving the ARI signal
to zero. The ARI signal is also driven to zero with wheel spin
up. This is one of the aids for better control during crosswind
landings. With ARI operational during a nose-high rollout, and
with lateral stick held to lower the up-wind wing, ARI would add
rudder in a direction to drive the nose into the wind.
Control Surface Actuators - The muscle for the Eagle's rudders is not the conventional linear ram-type actuator normally seen on aircraft flight control systems. To minimize space requirements, a rotary actuator was designed which is an integral part of the rudder hinge line. Not only does it receive manual inputs, but also electrical inputs from the servo amplifier. These signals control an internal piston which deflects the rudder control surface upon CAS command. The rudder actuator is also load-limited so that as inflight loads increase the actuator deflection is decreased accordingly, reducing unwanted tail loads. In the event of hydraulic pressure failure to either actuator, bypass valves prevent oil from escaping into the return line. In this condition, the rotary actuator becomes a self-contained surface damper.
Problem Areas - Yaw CAS (as well as CAS in general) has an excellent record of reliability. Once "infant mortality" rids the computers of problem components, the next malfunction is pretty far down the road. During production flights at St. Louis, we have replaced two rate gyro packages - one had a loose connection in an individual gyro, the other was out of tolerance gradient-wise. No accelerometer sensors have been replaced to date.
Some difficulty may be experienced when yaw CAS is initially turned on-the first pedal application may result in a yaw shutdown. The reason for this involves the locking rings which hold the rudder actuator CAS rams at zero while CAS is disengaged. During yaw CAS turn-on, these rings must move to unlock the CAS ram. Sometimes the unlock ring of one rudder actuator lags the ring in the other actuator. As rudder pedal force electrical signals attempt to move the rudders, one rudder CAS ram encounters the not-fully unlocked ring. This results in a slight lag in one of the rudders. If this lag exceeds four degrees, the failure detection circuits will be triggered. The corrective action, in this case, is to reset yaw CAS and you should be back in business with normal operation.
Roll CAS (Figure 2) provides stability augmentation (or roll damping) as well as supplying the maneuvering capabilities to satisfy Level I lateral control requirements throughout the F-15 envelope (Mil Spec MIL-F-8785B defines Level I as "Flying qualities clearly adequate for the mission Flight Phase.) Short period roll oscillations and aerodynamic disturbances are sensed by dual roll rate gyros whose outputs are shaped and amplified and sent to the stabilator actuators, not to the ailerons as one would expect (there are no electrical inputs of any kind to the Eagle's ailerons). The stabilator control surfaces operate differentially to stop the unwanted roll disturbance and restore stable flight.
Electrical commands from lateral force inputs to the pilot's control stick force transducrr (located both forward and aft in the TF-15A) are first applied to a deadband circuit in order to desensitize the roll commands around the neutral point. The roll command is then switched as a function of Mach number. In this process, the larger gradient assures that the time-to-bank requirements are available at the lower speeds. The lower gradient reduces the roll/yaw coupling tendencies at Mach numbers in excess of 1.5.
Dual roll rate gyros measure aircraft response to a lateral control stick input. The roll channel of the roll/yaw computer then adds or subtracts differential stabilator deflections to assure the proper response. Roll CAS authority is a maximum of ±5 degrees differential stabilator relative to the position selected by the mechanical control system.
The roll CAS error is limited by functions of airspeed and angle of attack. The airspeed limit is employed so that excessive structural loads are not generated on the differential tail. The scheduling signal is derived by the Dynamic Pressure Sensor Unit of the Automatic Flight Control Set. This unit receives pitot and static pressures which drive dual potentiometers. The output is then shaped to provide the proper gain schedule as shown in Figure 3. The gain potentiometer excitation voltages are switched through a lag network upon roll CAS engagement to reduce transients.
Additional roll CAS limiting is required to reduce the roll/yaw coupling for negative angles of attack, and at large positive angles, to minimize the adverse yaw. This provides an equivalent to the "aileron washout" function of the mechanical control system which was discussed in an earlier article. The angle-of-attack limiting is subtracted from the airspeed schedule as indicated in Figure 4.
No roll CAS is desired at angles of attack above 20 degrees. This prevents the adding of pro-spin controls through uncommanded pilot-induced CAS inputs, and roll damper inputs, at the higher angles of attack. The angle-of-attack schedule is switched to a fixed reference at wheel spinup to assure that full roll CAS authority is available for adequate crosswind control during landing rollout.
Failure Detection - As in the yaw CAS, there are a number of ways in which the system can detect and react to system abnormalities -
• Like the yaw channel, the Roll CAS sensors must track each other by preset limits. When these limits have been exceeded by any sensor, or if a failure occurs in the roll CAS computation electronics, roll CAS shuts down.
• Roll CAS also shuts down, and remains down, if yaw CAS fails or is turned off by the pilot. This assures that no adverse roll/yaw coupling will occur.
• Since roll and pitch CAS share the stabilator servo actuators, roll CAS will shut down if the pitch CAS fails. However, if there has been no failure of the roll computation, and if the stabilator servo actuators are still operative, the roll CAS can be reset and will operate normally. This is a pretty good troubleshooting aid. If a failure of pitch and roll CAS occurs, but roll can be reset, it is unlikely that the stabilator actuators, or the wiring to the actuators, are at fault. In this case, check the pitch CAS computation circuitry and the sensors.
• The primary means of detecting failures within the roll and pitch CAS servo loops consists of monitoring the level of error of a differential pressure sensor (DPS) hydraulic ram (one in each stabilator actuator). As long as the servo signals are equal, the DPS error is zero and the system will operate normally. Failure of a servo valve, or electrical failure of an actuator LVDT, will drive one or the other DPS ram hardover. As a result, roll and pitch CAS will shut down. A lag network is employed to filter the DPS error signal being monitored, minimizing nuisance shutdown. With the DPS failure detection scheme just described, a fast-operating, high-authority CAS can be employed with an acceptable level of failure transients for hardover servo valves, or in the event an actuator LVDT output is lost.
This presentation of the F-15 Control Augmentation System will be continued in our next issue of the DIGEST as we take a look at the pitch CAS.