Category Airplane Stability and Control, Second Edition

Spring Tabs

Spring tabs overcome the main problem of flying tabs, which do not provide the pilot with control of the main surface at low speeds, as when taxiing. In spring tabs, the pilot’s linkage to the tab is also connected to the main surface through a spring. If the spring is quite stiff, good low-speed surface control results. At the same time, a portion of the pilot’s efforts goes into moving the main surface, increasing controller forces.

Spring tabs have the useful feature of decreasing control forces at high airspeeds, where control forces usually are too heavy, more than at low airspeeds. At low airspeeds, the spring that puts pilot effort into moving the main surface is stiff relative to the aerodynamic forces on the surface; the tab hardly deflects. The reverse happens at high airspeeds. At high airspeeds the spring that puts pilot effort into moving the main surface is relatively weak compared with aerodynamic forces. The spring gives under pilot load; the main surface moves little, but the spring gives, deflecting the tab, which moves the main surface without requiring pilot effort.

The earliest published references to spring tabs appeared as Royal Aircraft Establishment publications (Brown, 1941; Gates, 1941). NACA publications followed (Greenberg, 1944; Phillips, 1944). But the credit for devising a generalized control tab model that covers all possible variations (Figure 5.14) belongs to Orville R. Dunn (1949). The Dunn model uses three basic parameters to characterize spring tab variations, which include the geared tab, the flying tab, the linked tab, and the geared spring tab.

Although the derivation of pilot controller force equations for the different tab systems involve only statics and the virtual work principle, the manipulations required are surpris­ingly complex. As is typical for engineering papers prepared for publication, Dunn provides only bare outlines of equation derivations. Readers of the 1949 Dunn paper who want to derive his final equations should be prepared for some hard labor.

Dunn concluded that spring tabs can produce satisfactory pilot forces on subsonic transport-type airplanes weighing up to several million pounds. At the time of Dunn’s paper, spring tabs had indeed been used successfully on the Hawker Tempest, the Vultee Vengeance rudder, all axes of the Canberra, the rudder and elevator of the Curtiss C-46 Com­mando, the Republic XF-12, and the very large Convair B-36 bomber. They also would be used later on the Boeing B-52 Stratofortress. Dunn’s account of the DC-6 development tells of rapid, almost overnight, linkage adjustments during flight testing. The major concerns in spring tab applications are careful design and maintenance to minimize control system static friction and looseness in the linkages.

The B-19 experience encouraged Douglas engineers to use spring tabs for many years afterwards. Both the large C-124 and C-133 military transports were so equipped. The

Spring Tabs

DC-6, -7, -8, and -9 commercial transports all have some form of spring tab controls, the DC-8 on the elevator and the DC-9 on all main surfaces, right up to the latest MD-90 version. In that case, the switch was made to a powered elevator to avoid increasing horizontal tail size to accommodate the airplane’s stretch. A powered elevator avoids tab losses and effective tail area reductions because tabs move in opposition to elevator travel.

The Douglas DC-8 and -9 elevator control tabs are actually linked tabs, in which pilot effort is shared between the tab and the elevator. This gives the pilot control over the elevator when on the ground. The DC-8 and -9 elevator linked tabs are inboard and rather small. The inboard linked tabs are augmented by outboard geared tabs, which increase the flutter margin over single large linked tabs. The DC-9 elevator controls are hybrid in that hydraulic power comes in when the link tab’s deflection exceeds 10 degrees. Spring tabs serve a backup purpose on the fully powered DC-8 ailerons and rudder and on the DC-9 rudder. The tabs are unlocked automatically and used for control when hydraulic system pressure fails. The same tab backup system is used for the Boeing 727 elevator.

The spring tab design for the elevators of the Curtiss C-46 Commando was interesting for an ingenious linkage designed by Harold Otto Wendt. Elevator surfaces must be statically balanced about their hinge lines to avoid control surface flutter. Spring tabs should also be statically balanced about their own hinge lines. Spring tab balance weights and the spring mechanisms add to the elevator’s weight unbalance about its hinge line. Wendt’s C-46 spring tab linkage was designed to be largely ahead of the elevator hinge line, minimizing the amount of lead balance required to statically balance the elevator.

Spring tabs appear to be almost a lost art in today’s design rooms. Most large airplanes have hydraulic systems for landing gear retraction and other uses, so that hydraulically operated flight controls do not require the introduction of hydraulic subsystems. Further­more, modern hydraulic control surface actuators are quite reliable. Although spring tab design requires manipulation of only three basic parameters, designing spring tabs for a new airplane entails much more work for the stability and control engineer than specifying

parameters for hydraulic controls. Computer-aided design may provide spring tabs with a new future on airplanes that do not really need hydraulically powered controls.

The 1956 Wright Field Conference

When inertial coupling first appeared on the scene with the Douglas X-3 research airplane and then the F-100A Super Sabre, interest in the subject grew quickly among those responsible for fighter airplane stability and control. Although first-line fighter flight test results were classified either confidential or secret, the grapevine was hard at work, and information started to circulate on this new potential for uncontrolled maneuvers and structural failures.

U. S. Air Force and Naval engineerssaw the need for groupsgrappling with the unexpected inertial coupling problem to convene and exchange information for the common good. A closed-door, classified conference was therefore called at Wright Field for February 1956. Papers were invited from industry, NACA, and MIT. Because of the urgency and national importance of the subject, authors and attendees from industry were expected to give open accounts of their results, putting aside competitive considerations.

The list of speakers at the since-declassified Wright Field Conference, formally called “Wright Air Development Center Conference on Inertia Coupling of Aircraft,” included many of the important stability and control researchers and designers of fighter aircraft of that period:

They were: Robert Bratt and Charles DaRos of Douglas; Frederick Curtis, Mamoree Masaki, and Dewey Mancuso of Convair; John Gautraud, James Flanders, Thomas Parsons, and Lloyd Wilkie of MIT; Richard Heppe of Lockheed; Wayne Huff and Cecil Carter of Chance Vought; Henry Kelley, Hans Hinz, and Robert Kress of Grumman; Darrel Parke of McDonnell; Jerry Pavelka of Republic; Stanley Schmidt, Norman Bergrun, Robert Merrick, Leonard Sternfield, Joseph Weil, and Richard Day of NACA; and John Wykes of North American.

Charles Westbrook chaired the conference and edited the proceedings (Westbrook, 1956). The lively interest in inertial coupling brought no fewer than 184 conference attendees. This conference on a serious stability problem held in the halls of their chief customer brought out

The 1956 Wright Field Conference

Figure 8.5 Time history of a classic inertial coupling example. The YF-102 diverges to a negative angle of attack and left (negative) side-slip angle in a rapid roll. (From Weil, WADC Conf. 56WCLC-1041, 1956)

a certain defensiveness in the speakers from industry. No blanket criticism is now intended, since it is understandable that airplane designers should want to put their products in the best light. Still, the transcript shows statements such as these:

In all of these [roll] tests the airplane response has been normal to the pilot and safe from every flight standpoint.

… serious difficulty due to inertial coupling is not to be anticipated for the——— .

…. the generally satisfactory roll behavior of the airplane was most welcome….

These benign, reassuring words were accompanied by hair-raising simulation and flight records, in several cases, such as the Convair YF-102 (Figure 8.5).

Longitudinal Control for Recovery

Tactical airplanes are able to reach supermaneuvering angles of attack by low or even negative static longitudinal stability. Full nose-up control starts the pitchup; unstable or nose-up pitching moment keeps it going. Recovery requires a nose-down pitching moment that will overcome the unstable pitching moment and leave a margin for nose-downward angular acceleration.

A rule of thumb for recovery nose-down pitching moment has been proposed, based on simulation studies and practical fighter design (Mangold, 1991). A pitching acceleration of 0.3 radians per second squared is said to be adequate. This leaves a margin for inertial coupling due to rolling during the pitching maneuver. A related problem is the amount of longitudinal control power required for very unstable airplanes, not necessarily during su­permaneuvers. For that problem, Mangold correlates required pitching acceleration control with time to double amplitude.

The recovery control problem also has been attacked using the classical Gilruth approach (Nguyen and Foster, 1990). Satisfactory and unsatisfactory recovery flight characteristics are used to draw a criterion line in a plot of minimum available pitching moment coefficients with full-down control versus a moment of inertia and airplane size parameter. With only five flight data points, Nguyen and Foster call their criterion preliminary.

10.5 Concluding Remarks

Current tactical airplane maneuverability research spans all aspects of the stability and control field, from linearized transfer functions to unsteady aerodynamics and the complex, vortex-imbedded flows found at very high angles of attack. Further advances and new theories appear likely with the advent of thrust-vectoring and direct side and normal force control.


Directional Stability and Control in Ground Rolls

The modern light plane tricycle landing gear has main wheels behind the center of gravity and a steerable castering, or freely swivelling, nose wheel ahead of the center of gravity. This arrangement, invented and applied by Fred C. Weick (1936), put an end to the ground loop. The ground loop is a rapid yaw from the runway heading and a swerve off the runway. It is a problem for tail wheel landing gears, which were still used by some designers for many years after Weick’s invention.

Directional Stability and Control in Ground Rolls

Figure 14.8 Forces and moments acting in wing sections and rolling tires. (a) Wing section. (b) Tire, top view. (From Abzug, 1999).

Directional Stability and Control in Ground Rolls

Figure 14.9 Ground roll eigenvalues for Cessna 182 at 3 airspeeds. (From Abzug, 1999)

The physical mechanism by which the main wheels of a tricycle landing gear cre­ate yaw stability during ground roll is explained in Weick’s 1936 paper. However, it is possible to model mathematically the landing rollout process in the same way we model flight dynamics (Abzug, 1999). The model produces either eigenvalues or roots for ground rollout small perturbations or nonlinear equations suitable for 6-degree-of-freedom transient analysis.

The keys to mathematical modeling of ground rolls are models for the forces and mo­ments applied to the airframe by tires in contact with the ground and for landing gear oleo struts. Both are available in the literature from automotive and aviation research. There is an interesting analogy between the forces and moments acting on wing sections and on rolling tires, as shown in Figure 14.8. Tire lateral force exhibits a linear relationship, up to a stall, with tire lateral slip angle, a sort of tire lift curve. Tire lift curve slope with slip angle is used to generate tire stability derivatives, which are added to the normal airplane small-perturbation equations of motion to produce eigenvalues in ground roll. Figure 14.9 has calculated eigenvalues for a Cessna 182 rollout at three airspeeds, as func­tions of main gear longitudinal distance from the center of gravity. Positive eigenvalues, indicating directional instability, occur for main gear locations just ahead of the center of gravity.

Linearized ground roll analysis can be applied to large airplanes with complex wheel arrangements and power-steered nose wheels, with less assurance of meaningful results. In those cases, linearized analysis may show ground handling trends, but one should plan
on full nonlinear 6-degree-of-freedom analysis, including tire forces. An extended analysis and simulation of ground roll was made for the Navy/Boeing T-45 trainer by the NASA Langley Research Center (Chambers, 2000). With proper tire dynamic models and inclusion of aircraft roll attitude, a pilot-induced yaw oscillation was reproduced.

The Penalty of Wing Sweepback on Low Subsonic Airplanes

Extra vertical tail length is obtained in canard configurations with wing-tip – mounted vertical tails by using wing sweepback. While we have learned how to provide good stall characteristics and a stable pitching moment stall break on sweptback wings,

The Penalty of Wing Sweepback on Low Subsonic Airplanes

Figure 17.1 Drawings of the tail-last Beech Super King Air B200 (above) and canard Starship 1 (below). The two airplanes are of similar size and gross weight, but the B200’s vertical tail length is 40 percent greater than the Starship’s. (From Jane’s AH the World’s Aircraft, 1987-1988)

these come at a cost in wing twist, special airfoil sections, or stall control devices such as slats, fences, and slots. Thus, wing sweepback used on a canard configuration to im­prove directional stability and control brings cost and weight penalties relative to tail-last configurations.

Aeroelastic Effects on Static Longitudinal Stability

There had been several published studies of the effects of aeroelasticity on static longitudinal stability, going back to 1942. But the subject really came to wide attention with the appearance of the very advanced, and flexible, Boeing B-47 Stratojet, first flown in 1947. Richard B. Skoog of NACA reported on the details of the stability and control static aero­elastic effects on this airplane (Skoog, 1957) based on classified work done six years earlier.

Strangely, while some individual effects are large, Skoog found that the overall aeroelastic modification to longitudinal stability and control is small (Figure 19.6). Wing symmetric bending causesthe wing tipsto wash out at increasing anglesof attack. Thismovesthe airload

Aeroelastic Effects on Static Longitudinal Stability

Figure 19.3 Arrangement of ailerons, spoilers, and flaps on the Boeing 707 airplane. The outboard and inboard flap-type ailerons are manually controlled, with the help of internal aerodynamic balance and balancing or geared tabs (here called servo tabs). The spoilers are of the slot-lip variety, located just ahead of the flaps. (From Cook, The Road to the 707, 1991)

relatively inboard, resulting in a forward, or destabilizing, shift of the wing aerodynamic center. However, there is a net loss in lift at positive angles of attack, a reduction in the lift curve slope. This is stabilizing, increasing the relative effect of the tail lift.

Fuselage bending under tail aerodynamic loads is destabilizing. That is, for upward tail loads, the fuselage bends upward at the rear, decreasing the tail angle of attack and the restoring moment of the tail. However, this effect is largely canceled by the downward bending of the aft fuselage under its own weight and the weight of the tail assembly, at the lower airspeeds associated with higher angles of attack. Just as wing torsion leads to aileron reversal at a sufficiently high airspeed, so does vertical bending of the aft fuselage lead to elevator or longitudinal control reversal. In the case of elevator controls, stabilizer twist adds to the problem (Collar and Grinsted, 1942).

The basic static aeroelastic analysis methods used up to the time when finite-element methods were introduced was the method of influence coefficients. Early expositions of the

Aeroelastic Effects on Static Longitudinal Stability

Figure 19.4 Two airplanes with slot-lip spoiler lateral controls to minimize loss in control power at high airspeeds due to wing twist: the McDonnell-Douglas DC-10 (above) and the Lockheed 1011 (below). In each case small outboard flap-type ailerons are used only at low airspeeds. (From NASA TN D 8373 and TN D 8360, 1977)

Aeroelastic Effects on Static Longitudinal Stability

Figure 19.5 Effects ofMach number and dynamic pressure (q) on the effectiveness of three alternate aileron designs for the Boeing 2707 SST. The spoiler-slot-deflector is effective at all airspeeds, while the tip aileron reverses in effectiveness around a Mach number of 1.0. (From Perkins, Jour. of Aircraft, July-Aug. 1970)

Aeroelastic Effects on Static Longitudinal Stability

Figure 19.6 The overall effect of flexibility on static longitudinal stability ofthe Boeing B-47 airplane. The net effect is moderate, a forward neutral point shift of only 7 percent at a dynamic pressure of 500 pounds per square foot. (From Skoog, NACA Rept. 1298, 1957)

influence coefficient method were given in Pai and Sears (1949) and in a classified NACA Research Memorandum of 1950, written by Richard Skoog and Harvey H. Brown.

As early as 1954 an important relationship was stated between frequency response and static aeroelastic characteristics. If aeroelasticity were a branch of pure mathematics, this relationship would be stated as a theorem, in these terms:

Airplane frequency response at frequencies below the lowest structural bending or tor­sional modes should agree with calculated rigid-body transfer functions when quasi-static aeroelastic effects are included.

This relationship, proved experimentally with the Boeing B-47 (Cole, Brown, and Holleman, 1957), provides an important check on static aeroelastic methods. In the 1980s, this relationship provided the basis for a comparison of alternate quasi-static aeroelastic methods for the Northrop B-2 Stealth Bomber.

Empirical Approaches to Pilot-Induced Oscillations

Figure 21.1 is a time history of the pilot-induced oscillation that occurred during landing of the Space Shuttle Orbiter Enterprise in 1977. Pilot-induced oscillations (PIO), or airplane-pilot coupling (APC) incidents, in which pilot attempts at control create instability, are a natural subject for pilot-in-the-loop analysis and a major motivating factor for the method’s development. However, pilot-induced oscillations appeared long before advanced pilot-in-the-loop methods were in place. Engineers were obliged to improvise solutions empirically, so that airplane programs could proceed.

One cause of pilot-induced oscillations was apparent without much deep study If con­trol surface rate of movement is restricted for any reason, such as insufficient hydraulic

Empirical Approaches to Pilot-Induced Oscillations

Figure 21.1 Time history of pilot-induced oscillations that occurred during landing of the space shuttle Enterprise, on October 26, 1977. Time lags in the longitudinal control system are considered to have been the primary cause. (From Ashkenas, Hoh, and Teper, AIAA Paper 82-1607-CP, 1982)

fluid flow rate into actuation cylinders, the pilot is unable to reverse control motion quickly enough to stop an airplane motion, once started. A late correction drives the airplane too far in the reverse direction, requiring ever-increasing control motions. Describing func­tion analysis of rate limiting does indeed show destabilizing phase lag. Thus, one empir­ical design rule for pilot-induced oscillation avoidance is high available control surface rates.

In unpublished correspondence W. H. Phillips comments on other empirical findings on pilot-induced oscillations:

We found that very light control forces together with sensitive control were very likely to lead to pilot-induced oscillations. Viscous damping on the control stick was not the answer as this put lag in the response to control force as well as the recovery. What was needed was a large force in phase with deflection for rapid stick movements, which could be allowed to wash out quite rapidly. This could be obtained with a spring and dashpot in series. Grumman

called this a “sprashpot” and used it successfully in the feel system of the F-11F________________ The

negative Cha of flap-type controls causes the control force to fall off after the airplane responds.

An additional empirical approach to solving longitudinal pilot-induced oscillation prob­lems is the double bobweight system described in Chapter 5. An aft bobweight provides heavy stick forces to start a pitch maneuver, by applying pitching acceleration forces to the stick. Stick force falls off as the airplane responds.

Challenge of Stealth Aerodynamics

The invention of aircraft that are almost invisible to ground or surface-to-air – missile radars promises to be an effective defensive measure for reconnaissance and attack airplanes. This development has taken six paths so far, the first three of which are a distinct challenge to stability and control designers:

Faceted airframes replace the smooth aerodynamic shapes that produce at­tached flows and linear aerodynamics. Radar returns from faceted shapes, such as the Lockheed F-117A, are absent except for the instants when a facet faces the radar transmitter.

Parallel-line planforms have the same sweep angle on wing leading and trailing edges and on surface tips and sharp edges. Parallel-line planforms concentrate radar returns into narrow zones that are easily missed by search radars. This is the Northrop B-2 stealth method, augmented by special materials and buried engines.

Suppressed vertical tails are either shielded from radar by wing structure or eliminated altogether. The Lockheed F-22 has shielded vertical tails, the B-2 none at all.

Blended aerodynamics eliminate internal corners such as wing-fuselage inter­sections. Internal corners can act as radar corner reflectors. The Rockwell B-1 uses this technique to reduce its radar signature.

Buried engines and exhausts hide compressor fan blades and hot exhaust pipes from radar and infrared seekers.

Radar-absorbent materials are used, generally nonmetallic. This is a highly classified subject.

The challenges of faceted airframes, parallel-line planforms, and suppressed vertical tails to stability and control engineers are illustrated by current stealth airplanes.

Wright Controls

In the Wright brothers’ 1902 glider and their 1903 Flyer the pilot had a vertical lever for the left hand that was pulled back to increase foreplane incidence. The pilot lay on a cradle that shifted sideways on tracks to cause wing warp. To roll to the left the pilot decreased the incidence of the outer left wings and increased the incidence of the outer right wings. The rudder motion was mechanically connected to the wing warp mechanism to turn the nose left when the pilot wished to lower the left wing, and vice versa for lowering the right wing, thereby overcoming the adverse yaw due to wing warp.

When they began to fly sitting up in 1905, the Wrights retained the left-hand vertical lever for foreplane incidence but added a right-hand vertical lever for wing warp and rudder. They moved the new right-hand lever to the left for left wing down and forward for nose – left yaw. The right-hand lever was moved to the right for right wing down and aft for nose-right yaw. Turn coordination required the pilot to phase control motions, leading with yaw inputs. These unnatural control motions had to be learned and practiced on dual control machines or simple simulators. Bicyclists to the last, they never used their feet for control. They retained this scheme until 1909. Since wing warping involved considerable elastic deformation of the wing structure, they later changed the fore-and-aft motion of the right – hand lever to wing warp and mounted a new, short lever on its top for side-to-side movement to control the rudder. When the Wrights abandoned the all-moving foreplane array for an all-moving rear horizontal tail in 1911, the left-hand lever still controlled its incidence, but now reversed.

The Wrights’ patent was for mechanically linked roll and yaw controls. Other airplane builders, notably Curtiss, built airplanes with ailerons, rudders, and elevators, providing independent three-axis control. Curtiss and others asserted that the Wright machine now had independent three-axis control, but U. S. courts upheld the Wright patent against them. The courts maintained that the coupling of roll and yaw controls in the Curtiss machines existed in the mind of the aviator and was essential to the art of flying. Therefore, the Curtiss independent three-axis control infringed on the Wright patent!

Modern Light Twin Airplanes

The situation is different again in the case of the modern light twin airplanes. The first of these planes was the five-to-seven place Aero Commander 520, introduced by the Aero Design and Engineering Corporation of Culver City, California, in 1950. A year or so later Beech introduced its Model 50 Twin Bonanza, Piper its Model PA-23 Twin Stinson (later called Apache), and Cessna its Model 310 twin. These aircraft and their successors have a great deal of appeal to aviators who regularly fly on instrument flight plans into bad weather and those who want the extra safety of a second engine.

Yet by the early 1980s the safety records compiled by the modern light twins did not bear out this expectation. Writing in the AOPA Pilot of January 1983, Barry Schiff pointed out that the fatality rate following engine failure in light twins was four times that for engine failures in single-engine airplanes. It seems that relatively low-time private pilots were being trapped by the yaw and roll caused by the failure of one engine at low speeds and altitudes.

The Beech Model 95 Travelair and its higher power military derivative, the U. S. Army’s T-42A, are good examples of what could happen. After several fatal stall-spin accidents following power loss on one engine a courageous Army pilot made a series of T-42A stall tests, with symmetric and antisymmetric power. His report told of moderate wing drop in symmetric stalls, but of vicious behavior in stalls with one engine idling. The airplane would roll nearly inverted, clearly headed for a spin.

The response of the Federal Aviation Administration (FAA) to this generic light twin hazard was not to require design changes, but to warn pilots and to stress recognition and compensation for single-engine failure during training and flight tests for multiengine pilot ratings. Pilots are drilled to instantly recognize the failed engine by the mantra “Dead foot, dead engine.” Since accidents occur during the incessant single-engine drills in training, there is a special minimum airspeed for “intentionally rendering one engine inoperative in flight for pilot training.”

This is Vsse, the fourth of the special airspeeds the poor pilot has to memorize in order to legally operate multiengine airplanes. The others are Vmc or Vmca, the minimum airspeed for control with the critical engine’s propeller windmilling or feathered, the other delivering takeoff power; and Vxse and Vyse, the best angle of climb and rate of climb airspeeds with one engine inoperative. Vyse has its own marking on the airspeed dial, a blue line usually used as the landing approach airspeed under normal conditions. Evidently, if an engine fails on landing approach, one wants the airplane to be already at its best airspeed to climb away or to lose as little altitude as possible. The four special airspeeds for multiengine airplane operation are added to nine other special airspeeds (six if the airplane has no wing flaps or retractable landing gear) to be remembered.

In spite of the FAA’s apparent disinterest in obliging light twin builders to design safe single-engine behavior into their airplanes, there have been some attempts made in this direction. There is an FAA-approved design retrofit of vortex generators for the upper wing surfaces of some light twins. The installation reduces Vmca, the minimum airspeed for control with an engine out (Figure 4.3). Vortex generators are tiny (about 2 inches square)

Modern Light Twin Airplanes

Figure 4.3 Vortex generators fitted to the upper wing surface of a Piper PA-31-3 50 Chieftan light twin-engine airplane, to reduce minimum single-engine control speed Vmca. This installation of 43 generators on each wing was designed by Boundary Layer Research, Inc., of Everett, WA.

low-aspect ratio wings that stick out of a surface. The tip vortices from a spanwise row of generators set at angles of attack energize the surface’s boundary layer by mixing in with it high-energy air from the surrounding flow. The energized boundary layer tends to remain attached, avoiding separation or stall.

According to John G. Lee (1984), vortex generators were invented by “an introspective and rather unapproachable loner” named Hendrik Bruynes, who used eight vortex generators to correct separation from the walls of the diffuser in a new 18-foot United Aircraft Research Department wind tunnel. While Bruynes was named in the vortex generator patent, Lee credits Henry H. Hoadley with the key idea of reversing the angles of alternate generators. The Forty-Second Annual Report of the NACA, dated 1956, flatly credits H. D. Taylor of United Aircraft as the developer of vortex generators; no mention is made of either Bruynes or Hoadley.