Category Airplane Stability and Control, Second Edition

As airplanes evolved from stick and wire contraptions to awesome supersonic machines, the pilot at the center of it all has not changed. Desirable maximum and minimum levels of pilot stick, yoke, and rudder pedal control forces required to steer and maneuver are much the same – but the engineering solutions that bring these forces about have changed with the times.

5.1 Desirable Control Force Levels

In 1936 and 1937, NACA research pilots and engineers Melvin N. Gough, A. P Beard, and William H. McAvoy used an instrumented cockpit to establish maximum force levels for control sticks and wheels. In lateral control the maximums for one hand are 30 pounds applied at a stick grip and 80 pounds applied at the rim of a control wheel. In longitudinal control the maximums are 35 pounds for a stick and 50 pounds for a wheel. Lower forces are desirable and easily attainable with modern artificial feel systems.

The Federal Aviation Administration (FAA) allows higher forces for transport-category airplanes under FAR Part 25. Seventy-five pounds is allowed for temporary application. However, the data compilation for the handbook accompanying MIL-STD-1797, a current military document, shows that a little over 50 percent of male pilots and fewer than 5 percent of female pilots are capable of this force level.

Gough-Beard-McAvoy force levels are generally used as maximum limits for conven­tional stick, yoke, and rudder pedal controllers, but much lower control force levels are specified for artificial-feel systems and for side-stick controls operated by wrist and fore­arm motions.

Computational fluid dynamic methods apply the power of modern digital comput­ers to the problem of estimating stability and control from drawings at the design stage. Finite-element methods are one form of computational fluid dynamics. The great power of computational fluid dynamics is its ability to deal with arbitrarily shaped airplanes. Even advanced handbook methods such as the DATCOM can fail to represent a truly unusual design.

Computational aerodynamics are not exactly new, in that approximate methods for the calculation of aerodynamic loads on arbitrarily shaped wings in subsonic flow have been available for many years. However, before the advent of the digital computer the number of unknowns in the mathematical solutions for the flow had to be kept low. As pointed out by Sven G. Hedman, one of the inventors of the modern vortex lattice method, the number of unknowns was kept low in the predigital computer era by assumptions for the wing’s chordwise and spanwise load distributions. That early work was done by Falkner (1943), who also coined the term vortex lattice theory.

Assumed load distributions are not needed for modern finite-element methods using the digital computer. The earliest applications of modern finite-element methods to the calculation of aerodynamic forces and moments appears to have been made at the Boeing Company in about 1960. This was the discrete loading vortex lattice method, developed independently by Sven G. Hedman and P E. Rubbert. The development of the vortex lattice method is a classic case of research people all over the world contributing steps to a remarkably useful result. For a detailed history, see De Young (1976).

A major disadvantage of free-spinning wind-tunnel model testing is that the Frisbee-throw method of launching the model into the tunnel prevents study of real-life stalls and spin entries. Also, model size and Reynolds number are limited by the wind – tunnel size. Both of these problems with spin tunnels are partially eliminated for drop or radio-controlled models. For spin research, NASA drops models from helicopters and air­planes and operates radio-controlled dynamically scaled models that resemble superficially those flown by hobbyists.

For example, a 1 /12-scale drop model of a typical fighter airplane would weigh about 300 pounds. The NASA Langley helicopter drop model work is led by Charles E. Libbey. Both radio-controlled and helicopter drop model testing go on at the NASA Plum Tree test site, near Poquoson, Virginia.

9.2 Remotely Piloted Spin Model Testing

Perhaps the closest that one can get to testing a new airplane configuration for spin entry and recovery conditions in advance of the real thing is the remotely piloted drop model technique originated by Euclid C. Holleman and other NASA staff people at the Dryden Flight Research Center. A test pilot controls the drop model from a ground-based simulator­like isolated cockpit that has instrument displays run by telemetered data. The airplane’s flight control system is simulated on a ground computer and commands are up-linked to the model.

Considering that the setup resembles a space vehicle control center, design and operating expenses must be high compared with alternative spin model testing methods. However, the method efficiently produces good flight test data in a short time. Unplanned angles of attack and sideslip reached in remotely piloted tests of a 3/8-scale model of the McDonnell Douglas F-15 fighter might have resulted in loss of control for a less forgiving airplane.

On the negative side, pilots report that the absence of motion cues is a serious drawback for spin testing and might lead to overly conservative results. Holleman notes that before the actual drop tests pilots practice the flight plan on fixed-base ground simulators. Speeding up the pace for the fixed-base training sessions by a factor of 1.4 prepares the pilot for the actual flight.

It has been known for some time that landing approach path control by elevator or pitch adjustments does not work for low-aspect-ratio (stubby) straight or sweptback wings. This is due to the variation of drag with airspeed when the lift is equal to the gross weight in level flight conditions. We normally expect level-flight drag to increase rapidly with increas­ing airspeed, and so it does at cruising airspeeds. At cruising airspeeds height control by pitch attitude changes using the elevator is stable and effective. The throttle can be left fixed.

However, the level-flight drag for any airplane increases with decreasing airspeed near the stall, as a result of induced drag increases and flow separation at high angles of attack. As airspeed is reduced from cruising values level-flight drag reaches a minimum and then actually increases again as the airspeed is reduced still further. The airspeed at which level – flight drag, and thrust required to hold level flight, reach minimums was given the name “minimum drag speed” by Stefan Neumark in Britain (1953).

The increase in level-flight drag near the stall is accentuated for airplanes with low-aspect – ratio wings, leading to increases in minimum drag speed. The minimum drag speed for an airplane with a low-aspect-ratio wing can be well above the low approach airspeed desired for carrier landings. Thus, if an airplane with a low-aspect-ratio wing is on a stabilized descent at a low landing approach speed typically used for aircraft carriers and the pilot retrims the airplane to a higher angle of attack, reducing airspeed, the airplane will rise at first relative to the original path and then settle even faster. The flight path will become steeper, a counterintuitive result.

For landing approaches below the minimum drag speed, where increasing thrust is re­quired for decreasing airspeed in level flight, sometimes called “the back side of the thrust required curve,” pitch attitude control by the elevator is unsatisfactory, even with the throttle used to control height. Thrust control by the pilot or an automatic system (the Navy’s APCS) to hold constant airspeed or angle of attack has been used to artificially create the normal variation of thrust required for level flight.

“Backside” carrier-based approach problems were first recognized about 1950 (Shields and Phelan, 1953). Pilots needed to use higher approach speeds forthe XF-88A andXF3H-1 airplanes than the standard rules of thumb based on stalling speed. Shields and Phelan proposed a fixed-throttle pitch-up test maneuver that is similar to a popup maneuver later adopted as one criterion for minimum carrier-approach speed. The first large-scale organized set of data on minimum approach airspeed behavior for jet airplanes was taken at the NACA Ames Aeronautical Laboratory (White, Schlaff, and Drinkwater, 1957). Carrier – type landing approaches were made with seven straight – and swept-wing jet airplanes, the FJ3, F7U-3, F9F-6, F4D, F-100A, F-94C, and the F-84F. The objectives of the 1957 Ames tests was to find the minimum “comfortable” approach airspeeds for carrier-type landings for these representative jet airplanes.

The reason most frequently given by the NACA Ames pilots for minimum approach airspeeds was inability to control precisely altitude or flight path at lower speeds. However, there was a surprising lack of correlation between the minimum comfortable approach airspeed and the Neumark minimum drag speed. For example, Ames pilots set the minimum comfortable approach airspeed for the Douglas F4D-1 Skyray at 121 knots, while the minimum drag speed is 152 knots. Similar results appeared with the North American F-100A Super Sabre, where a minimum approach airspeed of 145 knots was selected, as compared with the minimum drag speed of 150 knots (Figure 12.2). Clearly, some other factors than inability of the elevator or stabilizer to control height without reversal were critical.

Another set of carrier-approach tests (Bezanson, 1961) found that flight path control of the Vought F8U and Douglas F4D-1 airplanes at low landing approach speeds required use of the throttle and was not satisfactory by angle of attack or pitch control modulation alone. Bezanson found that with thrust modulation as the primary path controller the dynamic characteristics of the thrust control system became important, including such factors as throttle friction and breakout force, throttle sensitivity (pounds of thrust per inch of throttle movement), and thrust time lag following abrupt throttle movements.

In contrast to pure jet engines, turboprops are operated at high RPMs all the time. Thrust modulation is done by propeller pitch changes, with very small time lags. The poor engine dynamic behavior of pure jet engines, particularly engine thrust time lag at low power levels (Figure 12.3), kept U. S. Navy interest alive in turboprop combat airplanes long after the U. S. Air Force had switched to pure jets. For example, the Douglas/Navy turboprop A2D-1 Skyshark made its first flight in 1950, the same year as the start of production on the Boeing/Air Force B-47A six-jet bomber.

Blind flying is controlled flight without reference to the outside scene, more specif­ically, the horizon. Simply put, there is no way that an airplane can be made suitable for blind flight by aerodynamic design alone. A pilot must rely on some form of gyroscopic device to retain control, either as a panel instrument or as part of an automatic pilot.

The need for instrumentation comes from the effects of spiral instability, or aileron control friction on airplanes that are spirally stable, as discussed in Sec. 5. Even if strong spiral stability were to be built into a design, at the expense of other desirable flying qualities, a pilot could lose control over altitude unless trained to damp the phugoid mode (Chapter 18, Sec. 9) by reference to airspeed alone.

Spatial disorientation due to illusions can be prevented by reference to instruments. The FAA’s Aeronautical Information Manual (1999) lists no fewer than 14 flight illusions that have been identified, such as a Coriolis, graveyard spin, somatogravic, and inversion illusions.

A mathematical model of the earth’s atmosphere is needed for stability and control flight simulation and other computer programs. These programs typically use dimensionless stability derivatives in setting up equations of motion for flight dynamics studies, and stability augmenter and autopilot analyses.

Standard atmospheric mathematical models were published by NACA starting in 1932. A 1955 model covered altitudes up to 65,800 feet (ICAO, 1955). NACA, the U. S. Air Force, and the U. S. Weather Bureau extended that model to an altitude of 400,000 feet (ICAO, 1962). The AIAA publishes a guide to standard atmosphere models (1996). For all its utility, the standard atmospheric model is based on quite simple assumptions: The air is dry, it obeys the perfect gas law, and it is in hydrostatic equilibrium.

Standard atmosphere computer codes for stability and control computer programs nor­mally accept as inputs the airplane’s altitude and true speed at each computing time. A minimum set of outputs at each computing time would include atmospheric density, Mach number, dynamic pressure, and equivalent airspeed. Additional outputs that could be gen­erated are static pressure and calibrated airspeed.

The standard atmosphere FORTRAN computer code shown in Figure 18.10 represents one of the two methods used in stability and control programs. In this example, air den­sity (RHO) is curve-fitted with exponential functions. Four function fits give satisfactory accuracy over the entire range of -4,000 to 400,000 feet. Speed of sound (ASPE), from which the Mach number is calculated, requires eight curve-fitted linear equations. The al­ternate standard atmosphere coding is ordinary interpolation from stored tables of density and speed of sound.

It has become increasingly important to represent wind gusts, shears, downbursts, and vortex encounters in stability and control flight simulation. Early flight simulations relied on a very simplified approach in which an additional gust angle of attack or sideslip is simply added to the values calculated at each instant from the airplane’s motions in an inertial space reference.

The sounder approach, now in common use, is an inertially fixed model for the wind environment, including gusts, shears, downbursts, and vortices. A NASA downburst model uses the conservation of mass principle to calculate wind velocity at all points within a down – burst (Bray, 1984). A central core is surrounded by an annular mixing region and a region of outflow parallel to the ground (Figure 18.11). In the Bray inertially fixed wind environment, an airplane penetrates the wind model as it moves along its path, just as in reality.

Earlier ad hoc wind shear models were proposed by NASA for flight simulation. A boundary layer shear model represents a low-level temperature inversion overlaid by strong winds. Two additional shear models, with the colorful names of the Logan and Kennedy shears, represent meteorologists’ best estimates of conditions existing at those airports during specific airplane wind shear encounters.

In Bernard Etkin’s terminology, the Bray downburst model and the NASA shear models are usually used as point atmospheric models, in which variations of local wind velocity over the airplane’s dimensions are neglected. Otherwise stated, the airplane is assumed to be vanishingly small with respect to the wavelengths of all spectral components in the turbulent atmosphere. This assumption obviously fails for gust alleviation systems that depend on sensing devices that sample air turbulence ahead of the main structure.

Etkin (1972) provides a thorough study of the finite airplane case, in which local wind ve­locities vary over the airplane’s dimensions. The required mathematics are surprisingly com­plex because atmospheric turbulence is a random process, and only a statistical, probabilistic

SUBROUTINE ATMOS (ALT, VEL, R HO, AMACH, DYN, VEKT)

: NASA/USAF/USWB STANDARD ATMOSPHERE, -4,000 TO 400,000 FT. DIMENSION X(3)

X(3) = ALT

IF(-X(3)-35.E3) 300,300,310 300 Z1 = 342.5E2 + X(3)*4.3/35.

GO TO 500

310 IF(-X(3)-45.E3)320,320,400 320 Z1 = 299.5E2-.285*(-X(3)-35.E3)

GO TO 500

400 IF(-X(3)-60.E3) 405,405,410 405 Z1 =271 .E2-.1 2*(-X(3)-45.E3)

500 RHO = .002377*EXP(X(3)/Z1)

GO TO 490

410 IF(-X(3)-140.E3) 415,415,420

415 RHO = EXP(-4.67263E-5*(-X(3)-6.E4)-8.41 364)

GO TO 490

420 IF(-X(3)-240.E3) 425,425,430

425 RHO = EXP(-3.871 2E-5*(-X(3)-14.E4)-1 2.1 51 584)

GO TO 490

430 RHO = EXP(-5.18378E-5*(-X(3)-24.E4)-16.022785)

490 CONTINUE

IF(-X(3)-362.E2)600,600,700 600 ASPE = 1117.-149.*(-X(3)/362.E2)

GO TO 800

700 IF(-X(3)-66.E3) 710,710,720 710 ASPE = 968.

GO TO 800

720 IF(-X(3)-105.E3) 730,730,740

730 ASPE = 6.71 282E-4*(-X(3)-66.E3) + 968.

GO TO 800

740 IF(-X(3)-1 555.E2) 750,750,760 750 ASPE = 1.73941 E-3*(-X(3)-105.E3) + 994.1 8 GO TO 800

760 IF(-X(3)-1 72.E3) 770,770,775 770 ASPE = 1082.02 GO TO 800

775 IF(-X(3)-200.E3) 780,780,785 780 ASPE = -1.21 5714E-3*(-X(3)-1 72.E3) + 1082.02 GO TO 800

785 IF(-X(3)-2625.E2) 790,790,795 790 ASPE = -2.6 236 8E-3*(-X(3)-200.E3) + 1047.98 GO TO 800 795 ASPE = 884.

800 AMACH =VEL/ASPE DYN = (RHO/2.)*VEL* *2 VEKT = 1 7.1 861 21 6*SQRT(DYN)

RETURN

END

Figure 18.10 FORTRAN digital computer subroutine for the NASA/USAF/USWB standard atmo­sphere. Air density (RHO) and speed of sound (ASPE) are curve-fitted in altitude bands from -4,000 to 400,000 feet. The subroutine requires inputs of altitude (ALT) and true speed (VEL). The subroutine outputs density, Mach number (AMACH), dynamic pressure (DYN), and equivalent airspeed (VEKT). (From ACA Systems, Inc. FLIGHT program)

treatment can be made (Ribner, 1956). Local wind velocity is a random function of both space and time. It simplifies things to assume stationarity, homogeneity, isotropy, and Gaussian distributions. Experimental data exist that provide adequate turbulence models for both high altitudes and near the ground, where isotropy does not hold.

More exotic atmospheric disturbances are significant for operation at very high altitudes and at hypersonic speeds. Flight disturbances due to temperature shears experienced by the

Lockheed SR-71A and the North American XB-70 are discussed in Chapter 11, “High Mach Number Difficulties.” In anticipation of a National Aerospace Plane (NASP) that would fly hypersonically, NASA laboratories at Dryden, Marshall, and Langley and the McDonnell Douglas Houston operation collaborated on a sophisticated FORTRAN atmospheric model called the NASP Integrated Atmospheric Model (Schilling, Pickett, and Aubertin, 1993).

The model is suitable for real-time simulations as well as batch programs. It provides global coverage, from the ground to orbital altitudes. The NASP model features of partic­ular stability and control interest are the small-scale perturbations that include continuous turbulence and the “thermodynamic” perturbations of density, pressure, and temperature. Gusts and thermodynamic perturbations can be selected either in patches or as discrete upsets.

Isolated mountain ridges that lie at right angles to the direction of strong prevailing winds can generate a so-called mountain wave. Air over the top of such a ridge cascades in huge volumes onto the lower terrain to the leeward. It then seems to bounce off, rising, then falling, and then rising again in a series of diminishing waves, all parallel to the ridge line. A huge rotor, or horizontal vortex, forms to the leeward of the first bounce. The characteristic mountain wave structure is well known to glider pilots, since glider altitude records are set by maneuvering into the rising air of the first bounce. Glider pilots also know to avoid the rotor, whose lower edge generally just brushes ground level. The National Transportation Safety Board (NTSB) concluded that a rotor was a possible cause for at least one airline accident. This was United Airlines Flight 585, a Boeing 737 lost at Colorado Springs in 1991. Various jet aircraft encounters with a rotor are modeled by Spilman and Stengel (1995).

Vortex wakes left in the atmosphere by airplanes flying ahead can be a severe hazard, although the principles for avoiding vortex wakes are known. Vortex wake fields can be modeled for flight simulation (Johnson, Teper, and Rediess, 1974).

Airplane transfer-function denominator factors, or roots, govern airplane motions following initial disturbances. Stable roots, having negative real parts, lead to subsidence of oscillatory or aperiodic motions. The same is true for the denominators of closed-loop transfer functions, and early root locus work, such as the material in Dr. William Bollay’s 1950 Wright Brothers Lecture, dealt with roots, or poles, of the closed-loop denominator. Transfer function numerator factors are called zeros. A response survey to step inputs for a systematic variation in pole-zero combinations (Elgerd and Stephens, 1959) gives striking results, particularly for the case of two real poles and one real zero. Depending on whether the zero is between the poles or to the right, the step response appears either deadbeat or with a large overshoot.

Transfer-function zeros play an important role in the closed-loop responses of stability – augmentation systems. The details are too involved to go into here, but some examples can be touched upon. In the altitude or glide path loop in which errors are corrected by elevator or stabilizer control, a zero called 1 /Th1 can be in the right half of the root locus plane. This occurs on the back side of the power required curve, or at airspeeds below the minimum drag point. Loop closure drives a closed-loop real root into the right half-plane, with consequent divergence. An inner stability augmentation loop can correct this.

Another example is the complex zero associated with bank angle control by the ailerons. The Systems Technology, Inc., symbol for the undamped natural frequency associated with this zero is Шф. For values of Шф that exceed the Dutch roll undamped natural frequency rnd, loop closure excites the Dutch roll and closed-loop stability is degraded. A complete tutorial discussion of this problem, as well as the altitude control zero problem, is given by Duane T McRuer and Donald E. Johnston (1975).

An early pilot-induced oscillation experience with the Space Shuttle Orbiter ALT – FF5 landing approach tests (Powers, 1986) led to some concern about the shuttle’s com­bination of large airplane size and distance of the pilot ahead of the center of gravity, and initially reversed altitude response with elevon control deflections. On the shuttle, more than 2 seconds elapses before altitude starts to increase following a step nose-up elevon deflection (Phillips, 1979). F. A. Cleveland (1970) gives a value of 0.8 second for this de­lay parameter on the C-5A Galaxy and speculates that glide path control during landing approach would be unsatisfactory for time delays above 2 seconds.

Cleveland’s analysis of the effects of airplane size gives a linear growth in the altitude delay parameter with size. Time delay would reach 2 seconds on a C-5A scaled up about

2.4 times, requiring direct lift control or canard surfaces. Digital sampling delays in the pitch control circuit and elevon rate limiting are known to contribute to the Space Shuttle’s pilot-induced oscillation tendency. According to Robert J. Woodcock (1988)

extensive, continual training [for the Space Shuttle Orbiter] has resulted in very good landing performance (smooth landings with small dispersion, mostly on long runways, in smooth air).

The conclusion is that large airplane sizes might indeed lead to landing approach prob­lems related to delayed altitude responses following pitch control inputs and special design features to correct the problem.

An alternate criterion for the control response of very large airplanes is the generic re­sponse to abrupt control inputs, as shown in Figure 10.5. Grantham, Smith, Person, Meyer, and Tingas (1987) compared large airplane flying qualities with the generic pitch response requirements of MIL-F-8785C. Maximum effective time delays of 0.10 to 0.12 second are specified in MIL-F-8785C, based on tests of relatively small tactical airplanes. These requirements may be much too stringent for large airplanes in all but very high-gain oper­ations. Satisfactory, or Level 1, flying qualities are obtained in simulation of a Lockheed 1011-class transport for effective time delays of 0.15 second, suggesting even larger accept­able values for superjumbo jets. On the other hand, the effective longitudinal axis dynamics of the Boeing 777 compares with some fighters.

John Gibson (1995) suggests that the observed sluggish response of the Space Shuttle Orbiter is caused by poor control laws as much as airplane size. Increasing pitch rate overshoot in a pullup would improve the observed sluggish flight path angle response. This further suggests that the flexibility of fly-by-wire technology can provide the necessary quickness of response in large airplanes without requiring special features such as direct lift control or canards.

In the late 1980s a counterrevolution of sorts took place, a retreat from authori­tative military flying qualities specifications. A new document (1987), called the Military Standard, Flying Qualities of Piloted Vehicles, MIL-STD-1797 (USAF), merely identifies a format for specified flying qualities. Actual required numbers are filled into blanks through negotiations between the airplane’s designers and the procurement agency. As explained by Charles B. Westbrook, the idea was to let MIL spec users know that “we didn’t have it all nailed down, and that industry must use some judgement in making applications.”

A large handbook accompanies the Military Standard, giving guidance on blank filling and on application of the requirements. The handbook is limited in distribution because its “lessons learned” includes classified combat airplane characteristics. The Military Standard development for flying qualities is associated with Roger H. Hoh of Systems Technology, Inc., and with Westbrook, David J. Moorhouse, and the late Robert J. Woodcock, of Wright Field.

The demise of the authoritative MIL-F-8785 specification was part of a general trend away from rigid military specifications, with the intent of reducing extraneous and detailed management of industry by the government. Industry designers said in effect, “Get off our backs and let us give you a lighter, better, cheaper product” and “Quit asking for tons of reports demonstrating compliance with arcane requirements.” Some horror stories brought out by the industry people did seem to make the point. The Military Standard is in fact ideal for “skunk works” operations; their managers don’t like more than general directions.

However, the Military Standard seems to bring back the bad old days, the “straw man” requirements of the 1930s, established by pilots and engineers based on hunch and specific examples. It is as if the rational Gilruth method had never been invented. A justification of sorts for the counterrevolution is the tremendous flexibility provided stability and control designers with the new breed of digital flight control systems.

Literally, it is now possible to have an airplane with any sort of flying qualities that one can imagine. Tiny side sticks can replace conventional yoke or stick cockpit controls. Right or left stick or yoke controls no longer have to apply rolling moments to the airplane. Instead, bank angle, constant rolling velocity, or even heading change can now be the result. By casting off the bonds of the rigid MIL-F-8785 specification, a procuring agency can take advantage of radical, innovative control schemes proposed by contractors.

The ability of advanced flight control systems to provide any sort of flying qualities that can be imagined brought a cautionary note from W. H. Phillips, as follows:

The laws of nature have been very favorable to the designers of control systems for old – fashioned subsonic, manually-controlled airplanes. These systems have many desirable features that occur so readily that their importance was not realized until new types of electronic control systems were tried.

Don Berry, a senior engineer at the NASA Dryden Research Center, had similar views:

We have systems capable of providing a wide variety of control responses, but we are not sure what responses or modes are desirable.

A further step in the dismantling of “rational” Gilruth flying qualities specifications is the recent appearance of independent assessment boards, charged with managing the flying qualities (and some performance) levels of individual airplanes. Such a board, called the “Independent Assessment Team,” was formed for the Navy’s new T-45A trainer. Team members for the T-45A included the very senior, experienced engineers William Koven, I. Grant Hedrick, Joseph R. Chambers, and Jack E. Linden.

A very early application of plug ailerons was to the Northrop P-61 Black Widow, which went into production in 1943. The P-61 application illustrates the compromises that are needed at times when adapting a device tested in a wind tunnel to an actual airplane. The plug aileron is obviously intended to work only in the up position. However, it turned out not to be possible to have the P-61 plug ailerons come to a dead stop within the wing when retracting them from the up position. The only practical way to gear the P-61 plug ailerons to the cable control system attached to the wheel was by extreme differential. Full up-plug aileron extension on one side results in a slight amount of down-plug aileron angle on the other side. The down-plug aileron actually projects slightly from the bottom surface of the wing. Down-plug aileron angles are shielded from the airstream by a fairing that looks like a bump running spanwise.

Plug-type spoiler ailerons are subject to nonlinearities in the first part of their travel out of the wing. Negative pressures on the wing’s upper surface tend to suck the plugs out, causing control overbalance. Centering springs may be needed. There can be a small range of reversed aileron effectiveness if the flow remains attached to the wing’s upper surface behind the spoiler for small spoiler projections. Nonlinearities at small deflections in the P-61 plug ailerons were swamped out (as an afterthought) by small flap-type ailerons, called guide ailerons, at the wing tips.

Early flight and wind-tunnel tests of spoilers for lateral control disclosed an important design consideration, related to their chordwise location on the wing. Spoilers located about midchord are quite effective in a static sense but have noticeable lags. That is, for a forward-located spoiler, there is no lift or rolling moment change immediately after an abrupt up-spoiler deflection. Since airfoil circulation and lift are fixed by the Kutta trailing – edge condition, the lag is probably related to the time required for the flow perturbation at the forward-located spoiler to reach the wing trailing edge. Spoilers at aft locations, where flap-type ailerons are found, have no lag problems (Choi, Chang, and Ok, 2001).

Another spoiler characteristic was found in early tests that would have great significance when aileron reversal became a problem. Spoiler deflections produce far less wing section pitching moment for a given lift change than ordinary flap-type ailerons. The local section pitching moment produced by ailerons twists the wing in a direction to oppose the lift due

to the aileron. This is why spoilers are so common as lateral controls on high-aspect ratio wing airplanes, as discussed in Chapter 19, “The Elastic Airplane.”

Slot-lip spoiler ailerons are made by hinging the wing structure that forms the upper rear part of the slot on slotted landing flaps. Since a rear wing spar normally is found just ahead of the landing flaps, hinging slot-lip spoilers and installing hydraulic servos to operate them is straightforward. There is a gratifying amplification of slot-lip spoiler effectiveness when landing flaps are lowered. The landing flap slot is opened up when the slot-lip spoiler is deflected up, reducing the flap’s effectiveness on that side only and increasing rolling moment (Figure 5.11).