Category Airplane Stability and Control, Second Edition

World-Wide Flying Qualities Specifications

As mentioned earlier, the German air forces in World War II operated under a set of military flying qualities requirements related to the Gilruth set of 1943. The growth of civil aviation after the war led to a number of national and world-wide efforts to specify flying qualities requirements, in order to rationalize aircraft design and procurement in each country and the international licensing of civil aircraft. The goal of internationally agreed upon civil aircraft flying qualities standards is the responsibility of the International Civil Aviation Organization (ICAO), an arm of the United Nations. Annex 8 of the ICAO Standardsdealswith airworthiness, which includesadequate flying qualities(Stinton, 1996).

Standards have also been adopted by individual countries for both civil and military machines. An earlier section traced the evolution of U. S. flying qualities specifications for military aircraft. Similar evolutions took place all over the world. British military specifi­cations are in the UK DEF STAN publications. In particular, DEF-STAN 00-970, issued in 1983, is similar in style to MIL-F-8785C and provides much the same information (Cook, 1997).

British civil flying qualities requirements were embodied initially in the BCARs, or British Civil Airworthiness Requirements. European standards now apply, as found in the European Joint Aviation Requirements, or JARs, issued by the Joint Aviation Adminis­tration. The U. S. versions are the Federal Air Regulations, or FARs, parts 21, 23, 25, and 103 of which deal with airplanes. The wording of the stability and control airworthiness requirements of the FARs is similar to the Gilruth requirements of 1943, which were also concerned with minimum rather than optimum requirements.

Spoiler Opening Aerodynamics

Experimental or wind-tunnel studies of rapidly opening upper-wing surface spoil­ers show a momentary increase in lift, followed by a rapid decrease to a steady-state value that is lower than the initial value. At a wind speed of 39 feet per second, the initial increase is over in less than a half-second, and steady-state conditions appear in about 3 seconds (Yeung, Xu, and Gu, 1997). Results from the computational fluid dynamics method known as the discrete vortex method also predict the momentary increase in lift and associate it with a vortex shed from the spoiler upper edge in a direction that increases net airfoil circulation in the lifting direction. A subsequent shed vortex from the wing trailing edge in the opposite direction reduces circulation to the steady-state value. While suggestive, experimental flow visualization results do not exist that confirm this vortex model.

The Yeung, Xu, and Gu experiments show that providing small clearances between the spoiler lower edge and the wing upper surface reduces the momentary increase in lift following spoiler extension. This is consistent with a small shed vortex from the spoiler lower edge of opposite rotation to the vortex shed at the upper edge. A clearance between spoiler and wing surface of this type has also been used to reduce buffet.

The B-52 Elevator Also Has Limited Control Authority

The B-52’s elevator is as narrow in chord as is the rudder. It depends on help from an adjustable stabilizer for long-term trim and airspeed changes. As in the case of the vertical tail, the original Boeing design called for an all-moving horizontal tail, but this was abandoned because of doubts as to hydraulic actuator reliability.

The B-52’sadjustable stabilizer isdriven by two independent hydraulic motorsthrough an irreversible screw jack mechanism. One motor drives the jackscrew and the other the live nut on the driven screw thread (Figure 7.6). The control valve for each hydraulic motor is worked

The B-52 Elevator Also Has Limited Control Authority

Figure 7.5 B-52 Stratofortress in a crosswind landing attitude. The landing gears are pointed down

the runway while the airplane is yawed to the left, presumably into the relative wind. Crosswind landing gear reduces the need for rudder power. (From Loftin, NASA SP-468, 1985)

either by an electric motor or by a backup cable drive from the cockpit. The electric motors are controlled in turn by the usual push-button arrangement on the pilot’s control yoke.

With all of this redundancy, stabilizer adjustment failures can still occur, but the B-52 is landable in an emergency with elevator control alone, regardless of stabilizer position. Some center of gravity adjustment by fuel pumping is necessary for this to work.

Time Domain-Based Criteria

Time domain response specifications get around the need for equivalent systems. A standard time domain response form was used in the 1987 version of the U. S. flying

Time Domain-Based Criteria

Figure 10.6 Example pitch attitude bandwidth/phase delay criterion, with test results. (From Field and Rossitto, 1999).

Time Domain-Based Criteria

Figure 10.7 Pilot evaluation of pitch response using Gibson Nichols chart template. (From Blight 1996)

Time Domain-Based Criteria

Figure 10.8 Generic pitch rate response to abrupt control input. This type of transient response description has the advantage of applying to high-order stability-augmented as well as unaugmented airplanes. (From Mil Standard MIL-STD-1797, 1987)

qualities standard, MIL-STD-1797 (Figure 10.8). Other time domain response criteria have been proposed, as follows:

The C* Parameter L. G. Malcolm and H. N. Tobie originated the C* parameter, to blend normal acceleration and pitch rate responses to pitch control input. C* is actually a weighted, linear combination of the two responses, akin to the weighted performance indices used in optimization calculations.

The Time Response Parameter Some years later, C. R. Abrams enlarged on the C* parameter approach with a time response parameter that includes time delay in addition to the earlier normal acceleration and pitch rate terms.

Gibson Dropback Criterion This refers to the pitch attitude change following a commanded positive pulse in airplane angle of attack. Pitch attitude increases during the pulse. A pitch attitude decrease after the pulse ends is called a drop – back. A slight dropback is associated with fine tracking. A large or negative dropback (pitch overshoot) creates unsatisfactory pitch short-period behavior.

Special Time Response Boundaries Upper and lower boundaries for longitudi­nal response was a still later specification form, used widely for landing approach responses in addition to up-and-away flying. The space shuttle Orbiter’s longitu­dinal control response is governed by such boundaries (Figure 10.9), apparently established in simulation.

Gibson (2000) comments that the upper boundary in particular severely limits rapid acquisition of angle of attack change in response to pitch demand and was responsible for space shuttle touchdown problems. He says further:

Time Domain-Based Criteria

Figure 10.9 An example of a time response boundary. The pitch rate response to a step-type manip­ulator input must lie between the boundaries. Pitch rate response q is normalized by the steady-state value qss. This particular time response boundary applies to the space shuttle Orbiter. (From Mooij, AGARD LS 157, 1988)

The UK HOTOL project (a horizontal take off Shuttle equivalent) was studied at Warton.. .By designing to optimum piloted pitch response dynamics, i. e., with a rapid flight path response and hence considerable pitch rate overshoot, accurate automatic touch­down was easily achieved in simulation.

Further progress in understanding and improving longitudinal maneuverability has made use of closed-loop studies using the human pilot model (see Chapter 21).

Ultralight Airplane Pitch Stability

Conventional ideas about the need for longitudinal static stability are misleading in the case of ultralight airplanes. The reason is that, instead of the normal short – and long – period, or phugoid, modes of motion, four unfamiliar first-order modes may appear. For ex­ample, the Gossamer Condor’s center of gravity is aft of the neutral point, in order to unload somewhat the canard surface. This produces a positive or unstable value for the CMa deriva­tive. As a result, one of the four first-order modes is unstable. However, the corresponding divergence has a time constant of about 1,000 seconds, making it imperceptible to pilots.

Another way to explain the benign pitch behavior of ultralight airplanes flying at centers of gravity behind the neutral point is to consider their maneuvering stability. Maneuvering stability disappears at the maneuver point. The maneuver point of ultralight airplanes tends to be far aft of the neutral point because of high pitch and heave damping levels. For flight at centers of gravity behind the neutral point but ahead of the maneuver point, the machine would have no tendency to diverge unstably in pitch attitude at constant airspeed. Its unstable behavior would require a simultaneous loss of airspeed and nose-up pitch change in level flight, a process that is very slow.

To illustrate the concept of the maneuver point or maneuvering stability, consider an airplane with an unstable gradient in pitching moment with angle of attack, and suppose it to be disturbed nose-up with respect to its flight path. The unstable pitching moment gradient would tend to increase the size of the disturbance, but at the same time the increase of angle of attack would cause the flight path to curve upward if the speed is constant. The upward curvature of the flight path implies an angular velocity in pitch, which is resisted by the aerodynamic damping in pitch.

In the case of the Gossamer aircraft, the stabilizing effect of the pitch damping due to flight path curvature overwhelms the destabilizing gradient of pitching moment with angle of attack. The neutral point is 5 percent ahead of the center of gravity, but the maneuver point is four chord lengths behind the center of gravity, due to the large path curvature for a given angle of attack.

The Oblique or Skewed Wing

Another rotation-only variable-sweep concept was invented by the late Robert T. Jones at the NACA Ames Aeronautical Laboratory, around 1945 (Figure 16.3). This is the oblique or skewed wing, in which wing sweepback (and sweepforward) is achieved by rotating the wing at its center, sweeping one side back and the other side forward. With the oblique wing rotated back into symmetry, the configuration avoids the tip stalling and low-speed stability and control problems associated with ordinary wing sweepback. Jones’ invention seems to have paralleled other rotating-wing sweep concepts, those of Lachmann of Handley Page and Richard Vogt of Blohm and Voss. Jones expected an additional advantage for the oblique wing as compared with conventionally swept wings, that of higher supersonic lift-drag ratio.

Had the unorthodox oblique-winged configuration been proposed by someone without Jones’ immense prestige, it might have been dismissed at once. But, for one thing, Robert Jones was credited with the invention of wing sweepback to alleviate compressibility ef­fects during World War II, independently of the Germans. He also contributed largely to stability and control theory, in all-movable controls, lateral control, two-control airplanes, and in solutions of equations of motion. Like the Wright brothers, Edward Heinemann, and John Northrop, Jones was not university-trained. His considerable mathematics were self-taught.

With the wing rotated, the oblique-wing configuration is that rarest of heavier-than-air machines, one without bilateral or mirror-image left-right symmetry Birds, dragonflies, and our own flying creations all have bilateral symmetry, as we ourselves do. It seems obvious that the flying qualities of an oblique-winged airplane would be strange, if not dangerous. For one thing, pulling up the airplane’s nose to increase angle of attack would create inertial rolling and yawing moments, quite absent in symmetrical airplanes. These moments arise from pitching velocity and acceleration acting on a nonzero product of inertia term Ixy.

The Oblique or Skewed Wing

Figure 16.3 Robert T. Jones, ahead of his time in many areas of aeronautics. He was the inventor in the United States of wing sweepback and of the oblique-winged airplane. Jones contributed to stability and control theory in lateral control, in two-control airplanes, and in all-movable controls. (From Hansen, Engineer in Charge, 1987)

The effectiveness of trailing-edge flap-type controls is seriously reduced at large sweep or skew angles. Control deficiencies can be made up if the airplane carries conventional tail surfaces. Control problems are more critical if an oblique wing airplane is always operated in the skewed position, but this would obviate the need for rotating engine pods and vertical tail surfaces.

Wing torsional divergence on the sweptforward panel, discussed in Chapter 19, “The Elastic Airplane,” has been raised as an issue for the oblique wing. Jones quite early predicted that rigid-body roll freedom would tend to raise the divergence speed to safe values outside of the flight range. That is, when the leading or sweptforward panel starts to bend upward under high airloads, the lift on that panel would increase, causing a large rolling moment. Airplane roll response to that moment would alleviate the airload and the wing would be safe.

However, the case must be considered in which automatic roll control operates to hold the airplane at zero bank. If the control rolling moment that holds the zero-bank angle comes from a horizontal tail, the wing torsional divergence speed could be close to the body-clamped case. A free-free analysis that includes autopilot loops would seem to be needed. On the other hand, if control rolling moment comes from ailerons on the leading panel, the panel loads would be reduced, as in the case of free-body roll. This would raise torsional divergence speed above the body-clamped value.

Some detailed stability and control data on oblique wings were obtained in NASA Ames Research Center wind-tunnel tests and in a NASA-Navy funded study begun in 1984. The

The Oblique or Skewed Wing

Figure 16.4 Zero-sideslip variations of rolling moment and side force coefficients for an oblique wing tested on a model of the NASA-Vought F-8 research airplane. Sizable, nonlinear values appear for wing skew angles of 30 degrees and above. (From Kroo, AIAA Short Course Notes, 1992)


study was on the feasibility of converting NASA’s F-8 Digital Fly-By-Wire Airplane to an oblique wing configuration. A key problem surfaced in the unusual nonlinear variations at zero sideslip angle in side force, and rolling and yawing moments with angle of attack, at wing skew angles as low as 30 degrees (Figure 16.4). These are trim moments, which would have to be trimmed out by control surface deflections for normal, nonmaneuvering flight. The nonzero side force could be equilibrated by flying at a steady bank angle, or possibly by wing tilt with respect to the fuselage.

Other possibilities to deal with nonzero side forces, yawing, and rolling moments at zero sideslip include wing plan form adjustments, unsymmetrical tip shaping, wing pivot selection, antisymmetric wing twist, and variable tip dihedral (Kroo, 1992). One is left with the impression that the aerodynamic design of a practical oblique-wing airplane will be far more complex than for its swept-wing counterpart.

There have been a number of oblique-wing flight tests, starting with a test in the Langley Research Center’s Free Flight Wind Tunnel. R. T. Jones also built and successfully flew a

The Oblique or Skewed Wing

Figure 16.5 The R. T. Jones invention in flight, the Ames-Dryden AD-1 oblique-wing testbed, flying with its adjustable wing in the fully swept 60-degree position. (From Hallion, NASA SP-4303, 1984)

The Oblique or Skewed Wing

Figure 16.6 The Vickers-Armstrong “Swallow” variable-sweep concept, tested at NASA’s Langley Laboratory and found to be longitudinally unstable with wings unswept. (From Polhamus and Toll,

NASA TM 83121, 1981)

series of small free-flying oblique-winged model aircraft, culminating in a radio-controlled two-meter span model whose wing andtailplane skew angle could be systematically changed in flight. A considerable number of oblique-wing design studies followed. NASA eventually contracted to have a full-scale oblique-wing test airplane built for low-speed flight tests. This airplane, the AD-1, a single-engine jet, was flown successfully at the NASA Dryden Flight Research Center at Edwards, California (Figure 16.5). Ten degrees of bank angle on the AD-1 are required to cancel the side force produced by a 60-degree wing skew angle (Kroo, 1992).

The Dryden flight tests were followed by a NASA design research contract for an oblique­wing supersonic transport. The contract was awarded to Boeing, McDonnell-Douglas, and Kansas University, around 1992. The study revealed arrangement problems with that partic­ular arrangement. An all-wing version of the oblique-wing eliminates the need for hinging the wing to a fuselage, although engine pods and any vertical tails still require hinging. Another NASA design research contract to Stanford University is for a flying model of a 400-foot-span all-wing supersonic transport, operated as an oblique wing. Stability and control for the all-wing versions of the oblique wing are problematic because of the problem of nonzero side forces, rolling, and yawing moments in oblique cruising flight.

The Semirigid Approach to Wing Torsional Divergence

In the semirigid approach to wing torsional divergence and related problems a reference section of the wing is selected to represent the entire three-dimensional wing. This simplification works quite well for slender wings, that is, wings of high-aspect ratio.

Semirigid analyses of wing torsional divergence are given in a number of textbooks (for example, Duncan, 1943; Fung, 1955). Fung shows a wing section that rotates about a pivot and is acted upon by a lift load. The pivot represents the chordwise location in the section of the wing’s elastic axis, or location where lift loads will not produce twist. The lift load can be taken as acting through the section’s aerodynamic center. The aerodynamic center, near the section’s quarter-chord point, is the point about which section pitching moments are invariant with angle of attack (Figure 19.1).

The wing section will come to a static equilibrium angle at some angle of attack under the combined action of the lift load and a spring restraint about the pivot. The spring restraint represents the wing’s elastic stiffness. If the pivot, representing the elastic axis, is behind

The Semirigid Approach to Wing Torsional Divergence

Figure 19.1 Semirigid model forwingtorsional divergence. Thewingis replacedby atypical section, pivoted aboutapoint that represents the wing’s elastic axis. The spring represents elastic stiffness. Inthis illustration, the wing’s aerodynamic center, where the lift acts, is forward of the pivot point. Increasing airspeed eventually leads to a torsional divergence. The angle of attack a increases without limit. (From Fung, The Theory of Aeroelasticity, Dover, 1969)

the wing’s aerodynamic center, the equilibrium angle of attack increases with increasing airspeed, which gives higher wing lift loads.

For any eccentricity, or distance of the aerodynamic center ahead of the pivot, and for given spring constants and wing lift curve slopes, or variations of wing lift with angle of attack, there is an airspeed at which the semirigid model diverges. That is, the equilibrium solution fails. Twist angle and angle of attack increase without limit. This is the calculated wing torsional divergence speed.

Wing torsional divergence problems were encountered on the Republic F-84 and Northrop F-89 airplanes, both equipped with large tip tanks (Phillips, 1998). Fixed fins on the outside rear of the F-84’s tanks moved the wing’s aerodynamic center aft, eliminating the problem.

Linear Quadratic Gaussian Controllers

Linear quadratic Gaussian (LQG) controllers add to the linear quadratic (LQ) designs random disturbances and measurement errors. LQG designs are discussed at length in a 1986 text and a 1993 IEEE paper by Professor Robert F. Stengel. The form taken by

Linear Quadratic Gaussian Controllers

Figure 20.8 Various control system forms that can be represented with the structured linear quadratic regulator(LQG)method. (FromStengel, IEEE Trans. onSystems, Man, andCybernetics, © 1993 IEEE)

the discrete-time LQG optimal controller is

uk = CFyk* – CBxk,

where yk * is the desired value of an output vector and xk is the Kalman filter state estimate.

The LQG design approach is very flexible because of the number of parameters that can be chosen arbitrarily. At one extreme, a scalar one-input, one-output design can be produced. Measurement and control redundancies can be represented if measurement and control vector sizes exceed that of the state vector. Also, integral compensation and explicit model-following structures can be produced (Figure 20.8).

LQG designs are among the most advanced to be in use by stability-augmentation en­gineers, as this is written. Even more advanced control concepts continue to pour out of university and other research centers. The same 1993 paper by Stengel cited above provides a good survey of advanced control concepts, including expert systems, neural networks, and intentionally nonlinear controls.

20.3 Failed Applications of Optimal Control

The failure of optimal control methods to produce a satisfactory flight control system for the Grumman X-29A airplane was noted in Sec. 14. This failure is by no means an isolated event. Additional instances can be found in which optimal control methods in the hands of experienced engineers have failed to produce safe and satisfactory flight control systems. What has gone wrong? Several experts who have witnessed these failures discuss the problem:

Phillip R. Chandler and David W. Potts (1983), U. S. Air Force Flight Dynamics Laboratory “[T]he infinite bandwidth constant compensation elements which are required [for LQR] violate the very heart of the feedback problem. . . . LQR therefore is an elegant mathematical solution to a nonengi­neering problem SVT (Singular Value Theory) [Doyle, 1979] is a very crude

method of coping with uncertainty in the LQR or LQG procedure. It makes assumptions that are not valid for flight control_______________________________________ LQR with all its ramifica­

tions and refinements is totally unsuited for the flight control servomechanism problem.”

John C. Gibson (2000), formerly with English/Electric/British Aerospace

“[Robert J.] Woodcock told me that there have been several missile and aircraft projects in serious trouble due to the use of such [LQG] methods_______________________________________ While op­

timization methods are continually being improved, they cannot yet (and may never) guarantee a safe and satisfactory FCS [flight control system] design with­out the strictest guidance and detailed physical understanding of experienced control and handling qualities engineers. This is true for highly advanced and demanding types of aircraft. Every signal path must be clearly visible and eas­ily related to specific aerodynamic or inertial characteristics of the airframe. In simple aircraft without complexity, there is no advantage over straightforward engineering methods anyway.”

Michael V. Cook (1999, 2000), Senior Lecturer, Cranfield University “There exists an enormous wealth of published material describing the application of so-called, ‘modern control methods’ to the design of flight control systems for piloted aeroplanes. It is also evident, with the exception of a very small number of recent applications, that there is a conspicuous lack of enthusiasm on the part of the airframe manufacturers to adopt this design technology, especially for the design of command and stability augmentation systems for piloted airplanes. Having an industrial background I am well aware of the many reasons why modern control has not been taken onboard seriously by the manufacturers – academic control specialists don’t share my view, and in many cases probably don’t even understand it!… I know that my views are shared by the control

people in——- who, in private are not at all complimentary about the academic

control specialists in the UK. I am also aware that the Boeing view is similar to

that of—— . I’ve seen some appallingly bad control systems design theses (not

from Cranfield).”

Steven Osder (2000), Osder Associates, Arizona “We [Osder and Dunstan Graham] used to lament the absurdity of papers [on robustness theory] that were filling the journals and we amused each other by citing specific examples of such departures from reason and logic_____________________ At the [Boeing] helicopter com­

pany, we took each of those University of – [robust flight control] designs and tested them against more complete [nonlinear] models of the [Apache] aircraft.

In every case, these robust flight control designs always fell out of the sky. In one case [which used eigenstructure assignment], even testing against a lin­ear model, but with only a 10 percent variation in a single B [control] matrix term, our simulations resulted in a crash.”

Duane T. McRuer (2001), Chairman, Systems Technology, Inc. “At STI we have spent an enormous amount of time and effort searching for ways to make optimal control practical – at least 20 major reports and papers, with some tremendously capable folk (e. g., Dick Whitbeck, Greg Hofmann, Bob Stapleford, Peter Thompson, et al.). Our focus has been on finding performance indices, special schemes, etc., to make optimal control solutions jibe with good

design practice__ We have just never been happy with the results for stability

augmentation design.”

In the light of the foregoing comments, a design case (Ward, 1996) in which an LQG design for a pitch stability augmentation system was used only as a guideline for a more conventional approach suggests a reasonable use for optimal control techniques. The concept of using LQR optimal control synthesis as a guide or in conjunction with classical methods is also developed by Blight (1996). Blight also comments that LQR methods should be used only on “control problems that actually require modern multivariable methods for their solution.” For example, Blight recommends ordinary gain scheduling instead of attempting to design a single robust linear control law for all flight conditions.


After raising student enthusiasm by a particularly inspiring airplane stability and control lecture, Professor Otto Koppen would restore perspective by saying, “Remember, airplanes are not built to demonstrate stability and control, but to carry things from one place to another.” Perhaps Koppen went too far, because history has shown over and over again that neglect of stability and control fundamentals has brought otherwise excellent aircraft projects down, sometimes literally. Every aspiring airplane builder sees the need intuitively for sturdy structures and adequate propulsive power. But badly located centers of gravity and inadequate rudder area for spin recovery, for example, are subtleties that can be missed easily, and have been missed repeatedly.

Before the gas turbine age, much of the art of stability and control design was devoted to making airplanes that flew themselves for minutes at a time in calm air, and responded gracefully to the hands and feet of the pilot when changes in course or altitude were required. These virtues were called flying qualities. They were codified for the first time by the National Advisory Committee for Aeronautics, the NACA, in 1943. Military procurement specifications based on NACA’s work followed two years later.

When gas turbine power arrived, considerations of fuel economy drove airplanes into the stratosphere and increased power made transonic flight possible. Satisfactory flying qualities no longer could be achieved by a combination of airplane geometry and restric­tions on center-of-gravity location. Artificial stability augmenters such as pitch and yaw dampers were required, together with Mach trim compensators, all-moving tailplanes, and irreversible surface position actuators. At roughly the same time, the Boeing B-47 and the Northrop B-49 and their successful stability augmenters marked the beginning of a new age.

Since then much of the art and science that connected airplane geometry to good low – altitude flying qualities have begun to be lost to a new generation of airplane designers and builders. The time has come to record the lore of earlier airplane designers for the benefit of the kit-built airplane movement, to say nothing of the survivors of the general-aviation industry. Accordingly, this book is an informal, popular survey of the art and science of airplane stability and control. As history, the growth of understanding of the subject is traced from the pre-Wright brothers’ days up to the present. But there is also the intention of preserving for future designers the hard-won experience of what works and what doesn’t. The purpose is not only to honor the scientists and engineers who invented airplane stability and control, but also to help a few future airplane designers along the path to success.

If this work has any unifying theme, it is the lag of stability and control practice behind currently available theory. Repeatedly, airplanes have been built with undesirable or even fatal stability and control characteristics out of simple ignorance of the possibility of using better designs. In only a few periods, such as the time of the first flights near the speed of sound, theoreticians, researchers, and airplane designers were all in the same boat, all learning together.

The second edition of this book brings the subject up to date by including recent de­velopments. We have also used the opportunity to react to the numerous reviews of the first edition and to the comments of readers. One theme found in many reviews was that the first edition had neglected important airplane stability and control work that took place outside of the United States. That was not intentional, but the second edition has given the authors a new opportunity to correct the problem. In that effort, we were greatly aided by the following correspondents and reviewers in Canada, Europe, and Asia: Michael V Cook, Dr. Bernard Etkin, Dr. Peter G. Hamel, Dr. John C. Gibson, Bill Gunston, Dr. Norohito Goto, Dr. Gareth D. Padfield, Miss A. Jean Ross, the late Dr. H. H. B. M. Thomas, and Dr. Jean-Claude L. Wanner.

The interesting history of airplane stability and control has not lacked for attention in the past. A number of distinguished authors have presented short airplane stability and control histories, as distinct from histories of general aeronautics. We acknowledge particularly the following accounts:

Progress in Dynamic Stability and Control Research, by William F. Milliken, Jr., in the September 1947 Journal of the Aeronautical Sciences.

Development of Airplane Stability and Control Technology, by Courtland D. Perkins, in the July-August 1970 Journal of Aircraft.

Eighty Years of Flight Control: Triumphs and Pitfalls of the Systems Approach, by Duane T McRuer and F. Dunstan Graham, in the July-August 1981 Journal of Guidance and Control.

Twenty-Five Years of Handling Qualities Research, by Irving L. Ashkenas, in the May 1984 Journal of Aircraft.

Flying Qualities from Early Airplanes to the Space Shuttle, by William H. Phillips, in the July-August 1989 Journal of Guidance, Control, and Dynamics.

Establishment of Design Requirements: Flying Qualities Specifications for Amer­ican Aircraft, 1918-1943, by Walter C. Vincenti, Chapter 3 of What Engineers Know and How They Know It, Johns Hopkins University Press, 1990.

Evolution of Airplane Stability and Control: A Designer’s Viewpoint, by Jan Roskam, in the May-June 1991 Journal of Guidance.

Recollections of Langley in the Forties, by W. Hewitt Phillips, in the Summer 1992 Journal of the American Aviation Historical Society.

Many active and retired contributors to the stability and control field were interviewed for this book; some provided valuable references and even more valuable advice to the authors. The authors wish to acknowledge particularly the generous help of a number of them. Perhaps foremost in this group was the late Charles B. Westbrook, a well-known stability and control figure. Westbrook helped with his broad knowledge of U. S. Air Force – sponsored research and came up with several obscure but useful documents. W. Hewitt Phillips, an important figure in the stability and control field, reviewed in detail several book chapters. His comments are quoted verbatim in a number of places. Phillips is now a Distinguished Research Associate at the NASA Langley Research Center.

We were fortunate to have detailed reviewsfrom two additional experts, William H. Cook, formerly of the Boeing Company, and Duane T. McRuer, chairman of Systems Technology, Inc. Their insights into important issues are used and also quoted verbatim in several places in the book. Drs. John C. Gibson, formerly of English Electric/British Aerospace, and Peter G. Hamel, director of the DVL Institute of Flight Research, Braunschweig, were helpful with historical and recent European developments, as were several other European and Canadian engineers.

Jean Anderson, head librarian of the Guggenheim Aeronautical Laboratory at the California Institute of Technology (GALCIT) guided the authorsthrough GALCIT’simpres – sive aeronautical collections. All National Advisory Committee for Aeronautics (NACA) documents are there, in microfiche. The GALCIT collections are now located at the Institute’s Fairchild Library, where the Technical Reference Librarian, Louisa C. Toot, has been most helpful. We were fortunate also to have free access to the extensive stability and control collections at Systems Technology, Inc., of Hawthorne, California. We thank STI’s chairman and president, Duane T. McRuer and R. Wade Allen, for this and for very helpful advice.

The engineering libraries of the University of California, Los Angeles, and of the University of Southern California were useful in this project. We acknowledge also the help of George Kirkman, the volunteer curator of the library of the Museum of Flying, in Santa Monica, California, and the NASA Archivist Lee D. Saegesser.

In addition to the European and Asian engineers noted previously, the following people generously answered our questions and in many cases loaned us documents that added materially to this work: Paul H. Anderson, James G. Batterson, James S. Bowman, Jr., Robert W. Bratt, Daniel P. Byrnes, C. Richard Cantrell, William H. Cook, Dr. Eugene E. Covert, Dr. Fred E. C. Culick, Sean G. Day, Orville R. Dunn, Karl S. Forsstrom, Richard G. Fuller, Ervin R. Heald, Robert K. Heffley, Dr. Harry J. Heimer, R. Richard Heppe, Bruce E. Jackson, Henry R. Jex, Juri Kalviste, Charles H. King, Jr., William Koven, David A. Lednicer, Dr. Paul B. MacCready, Robert H. Maskrey, Dr. Charles McCutchen, Duane T McRuer, Allen Y. Murakoshi, Albert F. Myers, Dr. Gawad Nagati, Stephen Osder, Robert O. Rahn, Dr. William P Rodden, Dr. Jan Roskam, Edward S. Rutowski, George S. Schairer, Roger D. Schaufele, Arno E. Schelhorn, Lawrence J. Schilling, Dr. Irving C. Statler, and Dr. Terrence A. Weisshaar.

Only a few of these reviewers saw the entire book in draft form, so the authors are responsible for any uncorrected errors and omissions.

This book is arranged only roughly in chronological order. Most of the chapters are thematic, dealing with a single subject over its entire history. References are grouped by chapters at the end of the book. These have been expanded to form an abbreviated or core air­plane stability and control bibliography. The rapid progress in computerized bibliographies makes anachronistic a really comprehensive airplane stability and control bibliography.

Malcolm J. Abzug E. Eugene Larrabee

Airplane Stability and Control, Second Edition

Power Effects on Stability and Control

The World War II period 1939-1945 coincided almost exactly with the appearance of power effects as a major stability and control problem. Grumman Navy fighters of that period illustrate the situation. World War II opened with the F4F Wildcat as the Navy’s first-line fighter and ended with the debut of the F8F Bearcat. The external dimensions of the two aircraft were almost identical, but the F8F’s engine was rated at 2,400 horsepower, compared to 1,350 horsepower for the F4F.

In unpublished correspondence, W Hewitt Phillips remarks that the appearance of power effects as a major stability and control problem was not entirely the result of growth in engine power:

these effects have been with us since World War I, but weren’t serious then because of the light control forces required to offset these effects, resulting from the low speeds and smaller size of these airplanes. The power effects in terms of thrust and moment coefficients were probably of the same order as in the case of the World War II fighters. These effects would have been somewhat reduced because of the short nose moment arm of these planes, and because of the lower lift coefficients due to the lack of high lift flaps.

The further growth in power and stability effects on military propeller-driven aircraft was of course interrupted by the advent of jets, with a different set of power effects on stability and control, generally of a minor nature. This chapter reviews the history of both propeller and jet power effects on stability and control. Although the days of high-powered propeller-driven military aircraft may be ended, their civil counterparts still exist, with a new set of stability and control problems.