Category Airplane Stability and Control, Second Edition

Internally Balanced Controls

Another control surface balance type that appeared about the same time as beveled controls was the internally balanced control. This control is called the Westland-Irving internal balance in Great Britain. Internally balanced controls are intended to replace the external aerodynamic balance, a source of wing drag because of the break in the wing contour. In the internally balanced control the surface area ahead of the hinge line is a shelf contained completely within the wing contour (Figure 5.12). Unless the wing is quite thick and has its maximum thickness far aft, mechanical clearance requires either that the shelf be made small, restricting the available amount of aerodynamic balance, or control surface throws be made small, restricting effectiveness.

By coincidence, internally balanced controls appeared about the same time as the NACA 65-, 66-, and 67-series airfoil sections. These are the laminar flow airfoils of the 1940s and 1950s. Internally balanced ailerons are natural partners of laminar flow airfoil sections, since aerodynamic balance is obtained without large drag-producing surface cutouts for the overhang. Not only that, but the 66 and 67 series have far aft locations of wing maximum thickness. This helps with the clearance problem of the shelf inside of the wing contour.

Internally Balanced Controls

Figure 5.12 The internally balanced control surface, used to reduce drag by eliminating the wing cutouts needed with overhang aerodynamic balances. Pressures at the upper and lower vent gaps are delivered to the sealed chamber, balancing the surface about its hinge. Pressure coefficients shown are for a 5-degree down-surface deflection. (From Toll, NACA Rept. 868, 1947)

An internal balance modification that gets around the mechanical clearance problem on thin airfoils is the compound internal balance. The compound shelf is made in two, or even three, hinged sections. The forward edge of the forward shelf section is hinged to fixed airplane structure, such as the tail or wing rear spar (Figure 5.13). The first application of the compound internal balance appears to have been made by William H. Cook, on the Boeing B-47 Stratojet. Internally balanced elevators and the rudder of the Boeing B-52 have compound shelves on the inner sections of the control surfaces and simple shelves on the outer sections.

Compound internal balances continue to be used on Boeing jets, including the 707, 727, and 737 series. The 707 elevator is completely dependent on its internal aerodynamic balance; there is no hydraulic boost. According to Cook, in an early Pan American 707, an inexperienced co-pilot became disoriented over Gander, Newfoundland, and put the airplane into a steep dive. The pilot, Waldo Lynch, had been aft chatting with passengers. He made it back to the cockpit and recovered the airplane, putting permanent set into the wings. In effect, this near-supersonic pullout proved out the 707’s manual elevator control. The 707’s internally balanced ailerons are supplemented by spoilers, as described in Chapter 19, “The Elastic Airplane.”

The later Boeing 727 used dual hydraulic control on all control surfaces, but internal aerodynamic balance lightens control forces in a manual reversion mode. An electrically driven adjustable stabilizer helps in manual reversion. At least one 727 lost all hydraulic power and made it back using manual reversion.

Internally balanced controls were used on a number of airplanes of the 1940s and 1950s. The famous North American P-51 Mustang had internally balanced ailerons, but they were unsealed, relying on small clearances at the front of the shelf to maintain a pressure

Internally Balanced Controls

Figure 5.13 Simple and two-element compound internal aerodynamic balances on the Boeing B-52 elevator. The compound balance segment is inboard. (From B-52 Training Manual, 1956)

differential across the shelf. The Curtiss XP-60 and Republic XF-12 both used internally balanced controls, not without operational problems on the part of the XP-60. Water col­lected on the seal, sometimes turning to ice.

The Phillips Inertial Coupling Technical Note

Electronic digital computerswere still yearsaway when Phillipsdid hisinertial cou­pling research. For numerical solutions that would attack the problem Phillips was obliged to simplify the equations with a series of ingenious mathematical steps. His successive transfor­mations led to inertial coupling stability boundaries derived by a simple quadratic equation.

For generality, Phillips nondimensionalized aircraft static stability parameters in terms of rolling frequency, or number of complete roll cycles per second. That is, the levels of both longitudinal and directional static stability or stiffness are characterized by their respective nonrolling natural frequencies, in short-period longitudinal and Dutch roll modes. These frequencies, expressed as cycles per second, are divided by the rolling frequency, as defined above, for the Phillips charts (Figure 8.2).

The remarkable but strong mathematical transformations added to the academic flavor of the charted results obscured the work’s significance to the hard-pressed stability and control engineers working in aircraft plants in the late 1940s, who should have paid more attention to Phillips’ results. Had a 1980s type digital computer been available to Phillips in 1947, permitting a few time histories of forthcoming fighter aircraft full-aileron rolls to

The Phillips Inertial Coupling Technical Note

Figure 8.1 The smoking gun – The XS-1 flight record that gave evidence of rapid oscillations in normal and lateral accelerations during steady rolling. The XS-1 drop model had an aileron wedge designed to make it roll steadily. (From Phillips, Jour. of theAmer. Aviation Historical Soc., Summer 1992)

The Phillips Inertial Coupling Technical Note

Figure 8.2 An example of the W. H. Phillips inertial coupling stability boundaries. (From Phillips,

NACATD 1627, 1948)

be calculated and presented, the airplane stability and control community would have taken notice.

Interesting background on Phillips’ inertial coupling work was contained in a 1994 letter from him. An excerpt from the letter reads:

In thinking about the subject lately, I have concluded that my approach was based on my training at MIT. In the courses that I took, particularly by Prof. Koppen, the derivations did not start with the complete equations of motion. The equations had already been divided into lateral and longitudinal groups and linearized. In Prof. Draper’s courses on instrumentation, much emphasis was placed on nondimensionalizing the results in terms of natural frequency. I did not read Bryan’s report, which starts from basic principles, for many years after that. If I had known the complete equations of motion, I might have been discouraged from attempting a solution.

While W. H. Phillips gave the first account of inertial coupling in the open literature, there seems to have been at least three other independent discoveries of inertial coupling. While working at the Boeing Company on the then-classified Bomarc missile, Roland J. White, Dunstan Graham, D. Murray, and R. C. Uddenberg found the problem and reported it in a Boeing Company document dated February 1948. At the Douglas Company’s El Segundo Division about the same time, Robert W. Bratt found inertial coupling in drop tests of a dummy Mark 7 bomb shape. A small amount of fin twist made the bomb spin. When the spin rate agreed with the bomb’s natural pitch frequency the spin went flat, or broadside to the wind.

Additional early work involving inertial coupling took place at the Cornell Aeronautical Laboratory in Buffalo, New York, by Donald W. Rhoads, John M. Schuler, and J. C. O’Hara. This was sponsored by the Structures Branch of the U. S. Air Force Wright Aircraft Laboratory, starting in 1949. Rhoads, Schuler, and O’Hara studied rolling pullouts, ma­neuvers that combine rolls and pullups. During the latter part of World War II vertical tail failures had occurred during rolling pullouts, as a result of large side-slip angles (Rhoads and Schuler, 1957).

Rhoads, Schuler, and O’Hara included inertial coupling terms in their study, among other refinements. Calculations of the critical peak side-slip angles agreed well with flight tests. However, their early numerical work, done at about the same time as the Phillips discovery, was for the Lockheed P-80 Shooting Star, whose inertial parameters are not much different than those of World War II airplanes. The P-80 has straight wings of moderately high aspect ratio and a fairly small value of the important inertial coupling parameter (Ix — Iy )/Iz. Inertial coupling was not prominent in the early stages of the Cornell Laboratory rolling pullout work, which actually extended over a period of five years. The stability and control community was not alerted.

The Phillips inertial coupling work, followed by flight occurrences of the phenomenon, led to a series of studies in Great Britain. W. J. G. Pinsker (1955, 1957, 1958) and H. H. B. M. Thomas (1960) were especially active.

Thus, the inertial coupling phenomenon, having been discovered in the late 1940s, was ignored by airplane designers until it was rediscovered in flight in the early 1950s. By 1956, the U. S. industry was roused enough to turn out for a conference on the subject held at Wright Field.

The Transfer Function Model for Unsteady Flow

Aerodynamicists familiar with the classical Bryan formulation of the perturbation equations of airplane motion expect to find aerodynamic forces and moments expanded in Taylor series. As an example, the yawing moment coefficient Cn is expanded as Cn = Cnp x в + Cnp x pb/2V + Cnr x rb/2V + Cns x 8 + ■ ■ ■. The series uses the first derivative only of the function (Cn) with respect to the independent variables, which are the vehicle’s state variables в, p, r, 8, etc. With this background, it is natural to treat unsteady flow effects by adding higher derivative terms to the expansion, such as Cn в x в.

Although mathematically sound, this approach has a serious flaw (Greenwell, 1998). Numerical values of higher order derivatives such as Cnp can be correct at only one oscillation frequency. Numerical values obtained in oscillating wind-tunnel rigs are correct at the frequency tested, but are in general invalid for the free or controlled angular motions of an airplane.

The solution of this problem is readily apparent to engineers trained in servomechanism theory. That is, treat aerodynamic force and moment as the result of dynamic processes much as hydraulic actuators and electrical networks are treated. The transfer function con­cept shown in Figure 10.2 is ideal for this application. Other modeling methods, such as Fourier function analysis, can produce equally valid results, but as Greenwell points out, the transfer function approach has the great advantage of being easily integrated into flight simu­lation computer codes (Abzug, 1997). Greenwell further proposes parallel transfer functions for applications at angles of attack that lead to separated flows and vortex bursting, each with its characteristic model. Transfer functions are not limited to first-order lag forms, but these have dominated the field so far. A first-order lag form adds one additional state to a state space aerodynamic model, as in the Goman and Khrabov example of Figure 10.11.

The Transfer Function Model for Unsteady Flow

Figure 10.11 Time-dependent mathematical model for aerodynamic force and moments proposed by M. G. Goman and A. N. Khrabov for the fully stalled regime encountered in supermaneuvers. Below, lift coefficient variation with angle of attack, using this model. (From Zgainov, AIAA Paper 93-4737,


The transfer function concept applied to modeling unsteady aerodynamics in simulations is typical of many developments in that it is difficult to establish priority. Greenwell credits Dr. Bernard Etkin as the originator of the concept, with his publication of a 1956 University of Toronto paper (UTIA Report 42). Early work on the concept also was done by Kenneth Rogers, Thomas Burkhart, J. Roy Richardson, Moti Karpel, William P Rodden, and R. Vepa.

A 100-state aeroservoelastic model ofthe GrummanX-29A forward-swept wing research airplane uses the transfer function model for unsteady aerodynamics. The transfer function model was also used with success at the DLR to model lift hysteresis at the stall for the Fairchild/Dornier Do 328 (Fischenberg, 1999) (see Chapter 14, Sec. 8.4).

10.3 The Inverse Problem

A requisite for linear analysis is a reference motion about which small perturbations occur. Generating reference motions in the case of fighter airplane supermaneuvers can be a particular problem. A planar reference maneuver, such as a Cobra-type snapup, may be generated with no particular difficulty from flight path equations (short-period dynamics suppressed) for a specific airplane. One would apply full nose-up aerodynamic or thruster control until the desired peak attitude or angle of attack is reached, followed by full nose- down control. This open-loop maneuver would yield a time history of airspeed, attitude, and angle of attack from which operating points could be selected for small-perturbation stability analysis.

Difficulties can be expected only if a maneuver path in inertial space is specified, rather than an open-loop time sequence of control or thruster angles. In that case, an inverse solution is required to determine the airplane’s velocity along the path and to be sure that the maneuver is possible in the sense that control limits are not exceeded. Again, specifying a planar reference trajectory for inverse solution presents no difficulty. The case is different for nonplanar maneuvers in that the geometry could become complex.

In principle, spatial sequences of six of the normal airframe states, the three position coordinates of the center of body axes and the three Euler angles, can define any airplane maneuver. A natural path set of coordinates has been proposed instead, particularly for nonplanar maneuvers (Myers, McRuer, and Johnston, 1987). The method is illustrated with the familiar yo-yo tactical maneuver. Natural path coordinates – tangent, normal, and binormal – are a familiar concept in classical mechanics.

Deep Stall

Deep stall requires a stable longitudinal trim point beyond the stall. If, in a stall, the combination of longitudinal trim and control is insufficient to nose the airplane down to an unstalled attitude, the condition is called a deep stall or locked-in deep stall. In some cases short of a locked-in stall, control power is so marginal that recovery takes place slowly or requires unusual measures, such as rolling or sideslipping the airplane or rocking the airplane in pitch. Deep stall was first encountered in-flight on a Handley Page Victor bomber in 1962. Hawker Siddeley’s Trident 1C, British Aircraft Corporation’s BAC 1-11, and the Soviet Tu 134 subsequently experienced deep stalls. The details of a BAC 1-11 deep-stall crash were widely disseminated, leading to a new series of wind-tunnel tests of airplanes then under development, such as the McDonnell Douglas DC-9. There were subsequent accidents in which deep stall was suspected, notably on the Boeing 727, a jet that resembles the BAC 1-11. The Canadair (Bombardier) Challenger CL600 also resembles the BAC 1-11. A CL600 and a variant were lost in deep-stall accidents.

There is an underlying cause for deep stall in airplanes with horizontal tails mounted on top of the vertical tail, in the T position. The wing trailing vortex system is normally rolled up into concentrated vortices by the time it reaches the horizontal tail. In unstalled flight the rolled-up vortices are generally behind the wing tips, quite a bit outboard of the horizontal tail span. This weakens the downwash at the horizontal tail, a source of nose-up pitching moment.

In airplanes prone to deep stall the outboard wing panels stall first when the angle of attack is increased. In effect, new wing tips are formed at the tips of the unstalled wing portion. The rolled-up trailing vortices are now quite close in span to the horizontal tail. If the horizontal tail is in the T location, at the top of the vertical tail, the rolled-up vortices at high angles of attack are roughly in the same plane as the horizontal tail, the position to exert a maximum downwash and nose-up pitching moment. The aerodynamic flow conditions for a deep stall on an airplane with a T-tail are illustrated in Figure 14.4.

NASA Ames 40- by 80-Foot Wind Tunnel tests of a full-scale Learjet, Inc., Model 23 executive jet provide a clear example of a deep stall at an aft center of gravity position (Soderman and Aiken, 1971). The airplane has a T-tail and a moderate wing sweep of 13 degrees. At the aft center of gravity position of 31.5 percent MAC, the data show a stable trim point at an angle of attack of 39 degrees, well beyond the stall. This is with full airplane nose-down stabilizer trim of 0.4 degree. With full 15-degree down-elevator added for recovery, the diving moment is insufficient to recover unstalled flight (Figure 14.5).

Deep Stall

Figure 14.4 Aerodynamic flows for deep stall for an airplane with a T-tail. Separated wing tip flow shifts the tip vortices inboard. Cores of the shifted tip vortices are in line with the T-tail, giving maximum downflow and a sharply increased down tail load.

The General Dynamics F-16 can also get into a deep stall (Figure 14.6). The F-16 has a special manual pitch override stick switch giving the pilot full tail travel, canceling roll and stability augmentation functions. This is to allow the pilot to rock out of a deep stall, in phase with residual pitch oscillations (Anderson, Enevoldson, and Nguyen, 1983). They report:

The instructions [flight manual] note that if no increase in attitude is discerned (with full [nose-up] pitch control), the pilot should wait 3 seconds and apply full reverse control. If the nose does not continue down with full forward stick, but reverses and starts up, full back stick must be applied to continue rocking the aircraft. The pitch oscillation has a period of approximately 3 seconds and the pilot is warned that rapid cycling of the control will be ineffective.

As many as four cycles are needed to break the F-16’s deep stall. Proper rock phasing is difficult if the airplane is in a roll oscillation, since severe rolling masks the airplane’s pitching motions. The ultimate fix for F-16 deep stall was a 25-percent increase in horizontal tail size, incorporated in all production airplanes (Chambers, 2000).

Deep Stall

Figure 14.5 Pitching moment coefficient variation with angle of attack for a full-scale Learjet Model 23, as tested in the NASA Ames 40- by 80-foot wind tunnel. At an aft center of gravity position of 31.5 percent MAC there is a stable trim point at an angle of attack of 39 degrees. This is with full nose-down trim. Circle symbols are flaps up, squares are flaps down. Addition of full down-elevator angle is insufficient to regain unstalled flight. (From Soderman and Aiken, NASA TN D-6573, 1971)

The McDonnell Douglas C-17 military cargo airplane also has a locked-in deep stall potential because of its T-tail. However, in contrast to the F-16 case, a sophisticated angle of attack limiter scheme prevents deep stalls from occurring. A control column pusher can prevent deep stall by overcoming the pilot pull force that is leading to a stall with an opposing push force. This approach was rejected because of reliability, particularly the possibility of a single point failure (Iloputaife, 1997).

The deep stall prevention system selected for the C-17 relies on a measured angle of attack providing an initial aural warning and shaking of the control column at a soft limit angle of attack. If the angle of attack continues to increase, or if airspeed drops at an excessive rate, the attack limiter system switches on, an angle of attack command system having a hard limit.

Interesting design features of the C-17 angle of attack limiter system are the array of six fuselage-mounted vane flow direction sensors and the provisions to switch the limiter in and out without causing undesirable transients or secondary stalls. The sensor array permits the system to operate correctly under sideslip conditions, which affect individual vane readings, and with redundancy in the face of vane failure or damage.

Deep Stall

Figure 14.6 Variation in pitching moment coefficient with angle of attack for the General Dynamics F-16A, with zero and maximum tail surface angles. Even with full nose-down control (+25 degrees) there is a stable or locked-in trim point at an angle of attack of 60 degrees. (From Nguyen, Ogburn, Gilbert, Kibler, Brown, and Deal, NASA TP-1538, 1979)

Canard Configuration Stall Characteristics

Canard aircraft are characteristically difficult to stall at all. The canard surface is generally designed to stall before the main, or aft, wing does, when the angle of attack is increased at a normal, gradual rate. When the canard does stall, with the main wing still unstalled, the airplane tends to pitch down, recovering normal flight. However, William H. Phillips comments that in the airplane pitch-down following canard stall, the canard surface’s angle of attack is increased again by the airplane’s angular velocity. This could delay recovery from an unstalled condition until the airplane has reached a steep nose-down attitude. It can be argued that an aft-tailed airplane also tends to recover automatically from a stall. On aft-tailed airplanes the horizontal tail, operating in the wing’s downwash, experiences a relative upload when the wing stalls. This is because the wing downwash drops off when the wing stalls.

The main concern in canard airplane stalls is the dynamic stall, entered at a high rate of angle of attack increase. Pitching momentum could carry the angle of attack up to the point where the main wing stalls, as well as the canard. In combination with unstable pitching moments from the fuselage, this could produce a total nose-up pitching moment that cannot be overcome by available canard loads. Wing trailing-edge surfaces that augment canard pitching moment control would be ineffective with the main wing stalled. Thus, a canard airplane’s main wing stall could produce deep stall conditions, in which a recovery to unstalled flight cannot be made by any forward controller motion (see Chapter 14). Deep stall at aft center-of-gravity positions and high power settings was identified in NASA tests of a tractor propeller canard configuration (Chambers, 1948).

The possibility of dynamic stalling on canard airplanes is minimized if the configuration is actually a three-surface case: main wing, canard, and aft horizontal tail. Examples of three – surface configurations are the Piaggio P180, the Sukhoi Su-27K, the DARPA/Grumman X-29A forward-swept research airplane, and the many three-surface airplanes designed by G. Lozino-Lozinsky, of MiG-25 fame. Even at extreme angles of attack that stall the main wing, the aft horizontal tail may be in a strong enough downwash field to remain unstalled, or it may be unstalled by nose-down incidence. With an unstalled aft horizontal tail, longitudinal control can be maintained.

Another way to minimize the possibility of dynamic stalling of canard airplanes is to operate them at centers of gravity far forward enough so that elevator power cannot produce high nose-up rotation rates. This amounts to restriction of the available center-of-gravity range and a reduction in the airplane’s utility.

Aileron-Reversal Flight Experiences

The World War II Japanese Zero fighter airplane had very low roll performance at high airspeeds, due to wing twist. U. S. combat pilots took advantage of this weakness. They avoided circling combat and established high-speed, single-pass techniques. At high airspeeds, the roll rates of the U. S. airplanes could not be followed by the Zeros, which were operating near their aileron-reversal speeds.

The role of aileron reversal due to torsional flexibility on missions of the Boeing B-47 Stratojet is mentioned in Chapter 3, “Flying Qualities Become a Science.” Boeing engineers attempted to deal with the roll reversal problem when designing the B-47 (Perkins, 1970). They knew there was a potential roll reversal problem since the B-47’s wing tips deflected some 35 feet between maximum positive and negative loads. Using the best approach known at the time, strip integration, torsional airloads were matched to stiffness along the wing span. This method predicted an aileron-reversal speed well above the design limit speed. Unfortu­nately, this approach didn’t take into account wing bending due to aileron loads. Wing bend­ing on long swept wings results in additional twist. The actual aileron-reversal speed turned out to be too low for low-altitude missions. Quoting from Perkins’ von Karman lecture:

A complete theoretical solution to the problem was undertaken at the same time [as the strip method application] and due to its complexity and the lack of computational help, arrived at the right answer two years after the B-47 first flew. A third approach to the problem was undertaken by a few Boeing experimentalists who put together a crude test involving a makeshift wind tunnel and a steel sparred balsa wood model that was set on a spindle in the tunnel with ailerons deflected and permitted freedom in roll. The tunnel speed was increased until the model’s rate of roll started to fall off and then actually reverse. This was the model’s aileron reversal speed and came quite close to predicting the full-scale experience. The test was too crude to be taken seriously and again results came too late to influence the design of the B-47.

According to William H. Cook, the B-47 not only had excessive wing torsional deflection due to aileron forces, but also slippage in the torsion box bolted joints. The wings would take a small permanent shape change after every turn. These problems led to a test of spoiler aileronsona B-47, although the production airplane wasbuilt with normal flap-type ailerons.

Decoupled Controls

Airplane stability augmentation must be rethought when designers choose to add direct normal and side force control surfaces. For example, with direct lift control through a fast-acting wing flap, pitch attitude can be controlled independently of the airplane’s flight path, and vice versa. The utility of such decoupled controls for tracking, defensive maneuvers, and for landing approaches is reviewed by David J. Moorhouse (1993).

20.4 Integrated Thrust Modulation and Vectoring

An airplane’s propulsion system can be integrated into a stability augmentation system that uses aerodynamic control surfaces. The total system would operate while the airplane remains under the control of the human pilot, qualifying as a stability-augmentation system rather than as an automatic flight control system.

For comparison, the previous coverage of propulsion systems in this book included:

Chapter 4 the effects of conventional, or fixed-configuration, propeller-, jet-, and rocket-propulsion systems on stability and control;

Chapter 10, Sec. 8 thrust vector control to augment aerodynamic surfaces in supermaneuvering;

Chapter 11, Secs. 14 and 15 propulsion effects on modes of motion and at hypersonic speeds;

Chapter 12, Sec. 1 carrier approach power compensation systems, for constant angle of attack approaches;

Chapter 20, Sec. 11 Propulsion-controlled aircraft, designed to be able to return for landing after complete failure of normal (aerodynamically implemented) control systems.

Depending on the number of engines under control, thrust modulation and vectoring systems can supply yawing, pitching, and rolling moments, as well as modulated direct forces along all three axes. Thus, thrust modulation and vectoring integrated into a stability – augmentation system can augment or replace the aerodynamic yawing, pitching, and rolling moments provided by aerodynamic surfaces. The situation is similar to aircraft like the Space Shuttle Orbiter, which carries both aerodynamic and thruster controls. However, in the context of stability augmentation, thrust modulation and vectoring would be used normally at the low airspeeds of approach and landing, rather than in space.

While in principle thrust modulation and vectoring can take the place of aerodynamic control surfaces at the low airspeed where the aerodynamic surfaces are least effective, it is reasonable to ask whether thrust stability-augmentation systems could satisfy flying qualities requirements. In a simulation program at DERA, Bedford (Steer, 2000), integrated thrust vector control was evaluated at low airspeeds on the baseline European Supersonic Commercial Transport (ESCT) design. The nozzles of all four wing-mounted jet engines were given both independent pitch and yaw deflections, providing yawing, pitching and rolling moments. Nozzle deflections were modeled as first-order lags. Conventional pitch rate, pitch attitude, velocity vector roll rate and sideslip command control structures were programmed.

Pitch control by thrust vectoring at approach airspeeds was as good as aerodynamic or elevon control, for a reason peculiar to the very low wing-aspect-ratio ESCT configuration. That is, the airplane has high induced drag at approach angles of attack, requiring large levels of thrust to maintain the glide path, thus making available large pitching moments with thrust deflection. Low airspeed roll and sideslip thrust vector control were positive and suitably damped but did not satisfy MIL-STD-1797A criteria.

Catching Up to the Wright Brothers

Two public demonstrations of perfectly controlled mechanical flight in 1908 by Wilbur Wright in France and by Orville Wright in the United States were clarion calls to the rest of the aeronautical community to catch up with and surpass their achievements. The airplane builders – Curtiss, Bleriot, Levavasseur, the Voisins, Farman, Bechereau, Esnault – Pelterie, and others – responded; by 1910 they flew faster and almost as well; by 1911 they flew better. However, even after these momentous achievements, neither the Wrights nor their competitors still had any real understanding of aerodynamic theory.

1.2 The Invention of Flap-Type Control Surfaces and Tabs

Flap-type control surfaces, in which a portion of the wing or tail surface is hinged to modify the surface’s overall lift, are at the heart of airplane control. Airplanes designed to fly at supersonic speeds often dispense with flap-type longitudinal controls, moving the entire horizontal surface. Also, some airplanes use spoiler-type lateral controls, in which a control element pops out of the wing’s upper surface to reduce lift on that side. Aside from these exceptions, flap-type controls have been the bread-and-butter for airplane control since a few years after the Wright brothers.

It was in 1908 that the aviation pioneer Glenn Curtiss made the first flight of his June Bug airplane, which was equipped with flap-type lateral controls. This was an early, if not the first, advance in lateral control beyond the Wright brothers’ wing warping. The Curtiss lateral controls were attached to the interplane struts between the biplane wings and were all-moving. Curtiss evidently saw them as lateral trim devices, since the wheel was connected to the rudder. The French called the flap-type lateral controls ailerons – little wings – and the name has persisted in the English language. The Germans call them querrudern, or lateral rudders.

The first true flap-type aileron control appears to have been on the French Farman biplane a year or two later. An aerodynamic theory for flap-type controls was needed, but it wasn’t until 1927 that Hermann Glauert (Figure 1.2) supplied this need. Control surface tabs are small movable surfaces at the trailing edge, or rear, of a flap-type control. Tabs generate aerodynamic pressures that operate with a long moment arm about the control surface hinge line. Tabs provide an effective way to deflect main control surfaces in a direction opposite to the deflection of the tab itself relative to the main surface.

Catching Up to the Wright Brothers

Figure 1.2 Hermann Glauert (1892-1934). In Glauert’s short career he made important airplane stability and control contributions, in control surface, downwash, airfoil, wing, and propeller theory, and in the equations of motion. (From Obit. Notices of Fellows of the Royal Soc., 1932-1935)

The tab concept is due to the prolific inventor Anton Flettner, who first applied it to steamboat rudders. One may still find references in the literature to “Flettners,” meaning tabs. Flettner received a basic German patent for the tab in 1922. This was for its application to aeronautics. Flettner’s patent includes a description of a spring tab device (see Chapter 5), which was promptly forgotten. Glauert’s aerodynamic theory for flap-type controls was extended to the tab case in 1928 by W. G. Perrin.

Direct-Thrust Moments in Yaw

For a multiengine airplane whose engines are mounted on the wings, when all engines are running and developing about the same power, there is no unbalanced yawing moment due to power. Failure of a wing-mounted engine of course sets up a thrust-caused yawing unbalance that must be counteracted by an equal and opposite aerodynamic yawing moment. The more engines on a multiengine airplane, the less effect will the failure of a single one have on yawing moments. Flight crews of Boeing B-29 and B-50 four-engine airplanes had the strange experience of losing engines during normal cruise flight and being unaware of it for many seconds. RPM for the dead engine would drop very little at first because of propeller windmilling. The directional stability of both airplanes was high enough to keep the ships close to course, initially.

Current Boeing Company design practice requires that a twin-engine jet transport be able to continue a climbout after takeoff, rudder-free, with one engine failed. This accounts for the generously sized vertical tails on the 737, 757, 767, and 777 models.

Safety Issues in Fly-by-Wire Control Systems

Although fully fly-by-wire flight control systems have become common on very fast or large airplanes, questions remain as to their safety. No matter what level of redundancy is provided, one can always imagine improbable situations in which all hydraulic or electrical systems are wiped out. Because of the very high-power requirements of hydraulic controls, their pumps are driven by the main engines. This makes necessary long high-pressure tubing runs between the engines and the control surfaces. The long high-pressure hydraulic lines are subject to breakage from fatigue; from wing, tail, and fuselage structural deflections; and from corrosion and maintenance operations.

The dangers of high-pressure hydraulic line breakage or leaking, with drainage of the system, could be avoided at some cost in weight and complexity with standby emergency electrically driven hydraulic pumps located at each control surface. An additional safety issue is hydraulic fluid contamination. Precision high-pressure hydraulic pumps, valves, and actuators are sensitive to hydraulic fluid contamination.

In view of rare but possible multiple hydraulic and electrical system failures, not to mention sabotage, midair collisions, and incorrect maintenance, how far should one go in providing some form of last-ditch backup manual control? Should airplanes in passenger service have last-ditch manual control system reversion? If so, how will that be accomplished with side-stick controllers?

In the early days of hydraulically operated controls and relatively small airplanes the answer was easy. For example, the 307 Stratoliner experience and other hydraulic power problems on the XB-47 led Boeing to provide automatic reversion to direct pilot control following loss in hydraulic pressure on the production B-47 airplanes. Follow-up trim tabs geared to the artificial feel system minimized trim change when the hydraulic system was cut out. Also, when hydraulic power was lost, spring tabs were unlocked from neutral.

Manual reversion saved at least one Boeing 727 when all hydraulic power was lost, and a United Airlines Boeing 720 made a safe landing without electrical power. The last-ditch safety issue is less easily addressed for commercial airplanes of the Boeing 747 class and any larger superjumbos that may be built. Both Lockheed L1011 and Boeing 747 jumbos lost three out of their four hydraulic systems in flight. The L1011 had a fan hub failure; the 747 flew into San Francisco approach lights. A rear bulkhead failure in Japan wiped out all four hydraulic systems of another 747, causing the loss of the airplane.

In another such incident the crew, headed by Delta Airlines Captain Jack McMahan, was able to save a Lockheed 1011 in 1977 when the left elevator jammed full up, apparently dur­ing flight control check prior to takeoff at San Diego (McMahan, 1983). There is no cockpit indicator for this type of failure on the 1011, and the ground crew did not notice the prob­lem. McMahan controlled the airplane with differential thrust to a landing at Los Angeles. This incident was a focus of a 1982 NASA Langley workshop on restructurable controls.

Workshop attendees discussed the possible roles of real-time parameter identification and rapid control system redesign as a solution for control failures.

Thus, although fully mechanical systems can also fail in many ways, such as cable misrig or breakage, jammed bellcranks, and missing bolts, questions remain as to the safety of modern fly-by-wire control systems. The 1977 Lockheed 1011 incident, a complete loss in hydraulic power in a DC-10 in 1989, and other complete control system losses led to the interesting research in propulsion-controlled aircraft described in Sec. 20.11.