A Flight Investigation of the Low-Speed Handling Qualities of a Tailless Delta-Wing Fighter Airplane

NASA MEMO 4-15-59/_ LO !

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NASA

MEMORANDUM

A FLIGHT INVESTIGATION OF THE LOW-SPEED HANDLING QUALITIES OF A TAILLESS DELTA-WING FIGHTER AIRPLANE By Maurice D. White and Robert C. Innis Ames Research Center Moffett Field, Calif.

NATIONAL

AERONAUTICS AND

SPACE

ADMINISTRATION

WASHINGTON

May 1959 Declassified September l, 1959 ii NATIONAL AERONAUTICS AND SPACE AEMINISTRATION _v_ORA__t_ 4-15-59A A FLIGHT INVESTIGATION OF THE LOW-SPEED HANDLING QUALITIES OF A TAYLI_SS DELTA-WING FIGHTER AIRPIANE* By Maurice D. White and Robert C. Innis Carrier landing-approach studies of a tailless delta-wing fighter airplane disclosed that approach speeds were limited by ability to control altitude and lateral-dlrectional characteristics. More detailed flight studies of the handling-qualltles characteristics of the airplane in the carrier-approach configuration documented a number of factors that contri- buted to the adverse comments on the lateral-directional characteristics.

These were: (1) the tendency of the airplane to roll around the highly inclined longitudinal axis, so that significant sideslip angles developed in the roll as a result only of kinematic effects; (2) reduction of the rolling response to the ailerons because of the large dihedral effect in conjunction with the kinematically developed sideslip angles; and (3) the onset of rudder lock at moderate angles of sideslip at the lowest speeds with wing tanks installed. The first two of the factors listed are inseparably identified with this type of configuration which is being considered for many of the newer designs and may, therefore, represent a problem which will be encountered frequently in the future. The results are of added significance in the demonstration of a typical situation in which extraneous factors occupy so much of the pilot's attention that his capability of coping with the problems of precise flight-path control is reduced, and he accordingly demands a greater speed margin above the stall to allow for airspeed fluctuations.

INTRODUCTION As a part of a general program being conducted at the Ames Research Center to investigate the landing-approach problems of hlgh-speed air- planes, flight tests were conducted on a number of airplanes, some results of which were reported in reference 1. One of the airplanes included in the study was the Douglas F4D-1, a tailless delta-wing fighter-type airplane, which was indicated to have flight characteristics in the *Title, Unclassified

landing approach that were different from those of most of the other

airplanes studied. In common with many of the other airplanes the

approach speed of the F4D-1 was reported to be limited primarily by

ability to control altitude. However, with the addition of external

underwing fuel tanks, substantially higher approach speeds were selected, an important limiting factor again being the ability to control altitude.

This occurred despite the fact that there were only slight differences

in the parameters that are usually assumedto affect ability to control

altitude, namely CLmax , drag variations with lift, thrust margins, etc.

A possible explanation for the difference in selected approach speed was

the fact that the lateral-directional characteristics of the airplane,

which were reported to be a secondary limitilg factor for the basic

airplane, were considered even worse when th._ wing tanks were added.

In order to determine quantitatively the factors that resulted in

this report by the pilots, a flight investigationwas conducted to docu-

ment the flying-qualities characteristics of the airplane in the landing-

approach configuration. The results of this investigation are presented

in this report.

NOTATION

mean aerodynamic chord drag coefficient

C D

lift coefficient

C L

pitching-moment coefficient about cen_er of gravity

Cm

cycles to damp to one-half amplitude C1/2 F control force, lb L lift-drag ratio D P static pressure, lb/sq ft ambient static pressure, lb/sq ft Po impact pressure, lb/sq ft qc ' time required for oscillation to damp to one-half amplitude, sec T1/2 time required for oscillation to double amplitude, sec T 2 V t_ sin B = V t_ _ for small _lues of V e 57.3 V indicated airspeed, knots true airspeed, ft/sec V t c_ angle of attack, deg sideslip angle, deg control-surface angle, deg angle of bank, deg rate of change of yaw angle, radians/sec CI ratio of air density at test altitude to that at sea level 191 57.3 ratio of bank-angle amplitude to equivalent side

Ivel

181 V t velocity amplitude for the oscillatory mode, deg ft/sec Subscripts a aileron e elevator r rudder AIRPLANE The Douglas F4D-I airplane is a tailless delta-wing jet-propelled fighter-type airplane. A three-view drawing and a photograph of the air- plane are shown in figure i, and pertinent physical characteristics are listed in table I. The airplane was equipped with two 300-gallon externally mounted (underwing) fuel tanks_ which were removed for the configuration described as the basic airplane. The engine is a Pratt and Whitney J57-PS-A, performance characteristics of which are presented in table I and figure 2.

Longitudinal and lateral control is obtained by actuation of power- boosted elevons on the wings together (longitudinal) or differentially (lateral). Typical variations of stick force and position with control deflection for an F4D-I airplane as extracted from reference 2 are shown in figures 3(a) and 3(b).

Directional control is obtained by actuation of two rudder segments.

The basic segment is moved by direct linkage to the rudder pedals. The variations of pedal force and position with rudder deflection from ref- erence 2 are shown in figure 3(c) with the yaw damper inoperative. The servo segment of rudder is slaved to the basic segment by an electrical sensing system and hydraulic drive. In the damper-on mode of operation the servo segment is also actuated by the following functions at the gearings indicated: (i) Yaw rate: 3° rudder per degree of _w per second (2) Aileron position: As shown in figtue 3(d) at speeds below about 200 knots (3) Side acceleration, as indicated by pendulum unit located behind pilot's headrest: gearing as shown in figure 3(e) INSTRUMENTATION Except as noted, the following items were obtained from transducer signals recording continuously on Consolidated oscillographs: Item TTansducer element Pressure sources on swivelling Airspeed and altitude pltot-static head on nose boom Swivelling vanes on nose boom Angles of attack and sideslip Stick, rudder-pedal, and control Control position transducers surface deflections Stick and rudder-pedal forces Strain-gage transducers Vertical and longitudinal NACA recording accelerometer accelerations Rates of roll, pitch, and yaw NACA re_ording turnmeters To calibrate the position error of the r_cording static-pressure source the airplane was flown by a ground installation for which the correct static pressure was simultaneously observed. The calibration curve obtained is shown in figure 4. The airspeed indicator which was connected to the service static-pressure source was calibrated by the same procedure, and the airspeeds reported by the pilot were corrected and rounded off for presentation in this report.

The elevator deflections presented are the average of the individual surfaces on the left- and right-wing panels.

TESTS, RESULTS, AND DISCUSSION Landing-Approach Investigations Flight tests to determine the minimum comfortable approach speeds in carrier-type landings were conducted at Crows Landing Auxiliary Landing Field as described in reference 1. Four Ames test pilots parti- cipated in the tests, and their selected approach speeds and reasons for limiting as extracted from reference 1 are shown in table II. A descrip- tion of the characteristics of the airplane that influenced the choice of approach speed is given here in more detail than in reference 1.

The primary reason for limiting the approach speed of the basic airplane (tanks off) was the ability to control altitude. Although all the pilots did not comment specifically on it, it was generally agreed that the lateral-directional stability and control characteristics of the airplane, which deteriorated with decreasing speed, were factors in the determination of the minimum comfortable approach speed. This was because the attention required to maintain lateral-directional control of the airplane diverted some pilot effort from the task of controlling flight- path angle precisely. With the addition of the external wing tanks, the pilots felt that the lateral-directional characteristics deteriorated considerably as indicated in table III. This factor then assumed about equal importance with the ability to control altitude in the pilots' determination of a minimum comfortable approach speed, and as a result, the pilots increased their approach speed by an average of about 9 knots.

The main objection to the lateral-directional characteristics lay in the low directional stability and the excessive adverse yaw coupled with the high dihedral effect. This behavior resulted in the lateral control producing considerable sideslip, but being ineffective as a roll control.

On the other hand, the application of rudder produced considerable roll in relation to the sideslip generated.

Lift-Drag Characteristics In figure 5 are shown the variations with lift coefficient of drag coefficient, angle of attack, and lift-drag ratio. Curves of drag against airspeed, as derived from the data of figure 5, are presented in figure 6. The data show no differences between the basic configuration and the tanks-on configuration that appear i_portant enough to account for the observed differences in approach speed.

Static-Longitudinal Stalility The variations of elevator angle and stick force with airspeed and with CL, as obtained from static measurements at varying speeds_ are shown in figure 7. The stick forces parallel the variations in elevator deflection because they are produced by a bungee which parallels the power driven, irreversible control system. This bungee is nonlinear, having different gradients for different control positions (fig. 3).

From the curves of elevator angle against CL, curves of C m against CL were computed with an assumed constant value of Cm5 e of -O.OOl_ per degree (fig. 8). The validity of the constant Cm_ e assumption is indi- cated by the good agreement of the above-mentioned curve of C m against C L with the curve of Cm against CL constructed from values of Cm_ as derived from the periods of short-period (sciL1_tions. Since the com- parison with period data was possible only for the tanks-on case, it was necessary to make the reasonable assumption that the same constant value of Cm5 e of -0.00145 was applicable to the basic configuration in determining the Cm versus CL curve of figure 8.

With tanks on, the variations of elevat(r angle or pitching-moment coefficient with lift coefficient (figs. 7 and 8) are reasonably smooth, but indicate a slight decrease in stability 1or a range of values of C L above about 0.5. For the basic configuratior, the smooth variation is broken by a small reversal in slope at lift coefficients above 0.75.

Neither of the aforementioned disturbances ir_ the stability curves bothered the pilots significantly because of the small magnitude of the decrease in stability in the tanks-on cases and becauEe the C L at which the slope reversal occurred with tanks off was well above the range of lift coefficients that would be used in an approach.

Dynamic Longitudinal Stability The results of limited tests of the short-period oscillation characteristics are shown in figure 9, where the period and damping variations with airspeed are plotted for the tanks-on configuration. A noteworthy characteristic shown by the data Js the low degree of damping, damping ratios being of the order of 0.2. The damping ratio, however, does not vary appreciablywith CL or airspeed. No data were obtained

for the basic configuration; however, in view of the similarity of the

other longitudinal characteristics over the operating range, it appears

likely that the dynamic stability characteristics of the basic configura-

tion would not be greatly different from those of the tanks-on configura-

tion shownin figure 9 which, in turn, showno outstanding feature that

would limit the approach speed.

Trim ChangeDue to Thrust

As noted in reference 3, one of the factors that has recently come

under scrutiny as possibly influencing the pilot's choice of approach

speed is the trim change due to throttle application. It has been noted

that airplanes differ in their initial response to throttle application

as a result of differences in trim changes due to thrust. In somecases

the added energy due to thrust is manifested primarily as a speed

increase; in other cases, there is predominantly a flight-path angle

increase, with little speed increase, or conceivably even a speed decrease.

Intermediate degrees of response between these extremes are also possible,

the ideal being a response with no speed changeswhatsoever. This partic-

ular characteristic would assumeincreased importance for airplanes of

the class of the F4D-1 in which landing approaches are madeon the "back

side" of the drag-velocity curve, where considerable throttle activity

would be required in making flight-path adjustments.

Unpublished data showthat the F4D-1 responses to throttle movement

were strong in speed changes and only moderate in flight-path angle

changes. While the pilots would have preferred that the trim changes

produce smaller speed variation, the over-all response characteristics

were considered acceptable because of a different compensating factor,

namely, the unusually large thrust margin available for maneuvering

(ref. 1).

Lateral-Control Characteristics Rapid control motions.- Time histories of the airplane response to abrupt aileron applications ihitiated from level flight at a speed of 125 knots are shown in figure l0 for the basic airplane and the airplane with tanks on. Data are shown both for the yaw damper on and off. The data show that for all the configurations the roll rate actually reverses after the first peak is attained, and is thereafter oscillatory about a level lower than that of the first peak. The resulting average or effec- tive roll rate is, of course, greatly reduced, but the amount of the reduction is not affected by the configuration.

The roll-rate reversals are attributed to rolling moments due to sidesllpangle, which have been shown to be very large for wings with swept leading edges operating at high angles of attack. Normally, the yawing moments that generate sideslip angles in this condition arise mainly from the adverse yaw characteristics of the ailerons. In the present instance, the sideslip angles result from another factor, namely, the tendency of the airplane to roll around its inclined longitudinal axis, so that sideslip would build up, at least initially, in accord with the relationship _ = m sin 2_. In figure ll, time histories of the function _ sin 2_, as computed from flight data, are plotted for compari- son with the recorded sideslip angles from figure 10. The agreement between the recorded sideslip angles and the sideslip angles computed as sin2_ is seen to be good for all the cases considered. Further veri- fication of the dominant role of the kinematics in generating sideslip dt included in is given by the time histories of the function /_ fig- ure ll. In the initial part of the time history, this function would indicate the main contribution of a yawing moment such as would be produced by aileron adverse yaw. It is apparent that, initially, the sideslip angles resulting from this source are much smaller than those arising from kinematic considerations, which indicates that the adverse yaw of the ailerons is of minor importance in defining the initial rolling responses to abrupt aileron control. It was necessary to confine these comparisons to the earlier stages of the maneuvers because the effects of side accelerations were not available to include in the comparisons.

It is of interest to note that the pilots could not detect from the air- plane motions the true source of the sideslips developed, but instead attributed it to aileron adverse yaw.

Slow control motion.- The test pilots re_orted that when the ailerons were moved slowly at low speeds the sideslip _u_gle tended to increase with little or no roll motion. This confirms the results of the abrupt aileron responses that the ailerons do produc._ some adverse yaw. Various combinations of aileron adverse yaw, static d2rectional stability, and dihedral effect could result in the observed z'esponses, which were, unfortunately, not documented in flight. As _Indicated by the variations of rudder deflection with steady sideslip an_.e in figure 12 and the derived curves of dSr/d _ in figure 13, the static directional stability of the airplane decreases with decreasing airspeed, so that the tendency of the airplane to yaw and not to roll in response to slow aileron movements would become more annoying at lower speeds as a result of this factor alone.

Steady Sideslips The variations of aileron, elevon, and rudder forces and deflections with steady sideslip are shown in figure ]2 for the basic airplane and for the airplane with tanks on. The variations of rudder angle with sideslip in figures 12 and 13 show a decreasing slope with decreasing speed, although the slopes indicate positive stability at speeds as low as 12_ knots. The rudder force gradients show similar trends except for the tanks-on case at 125 knots where a rudder force reversal is indicated for sideslip angles greater than about 6 ° . Sideslip angles of this order are attained in moderate aileron rolls (see fig. 10). The reason for this difference in rudder force variation at 129 knots between the basic configuration and the tanks-on configuration is not readily apparent.

It can only be surmised that the slightly lower rudder position gradient is sufficient to produce this effect, or that there is a difference in air flow over the tall due to the tanks that might account for it. In any case the pilots considered this characteristic dangerous enough that they were reluctant to extend the steady sideslip tests to higher sideslip angles for fear of producing a spin. It should be apparent that the need to maneuver the airplane with such considerations present would force the pilot to select higher approach speeds.

Lateral Stability In figure l_ are shown the variations with airspeed of the dynamic lateral stability parameters, period, damping, and roll-to-yaw ratio ]_]/]Ve]. Generally, there is a deterioration in damping (in terms of CI12) as speed is reduced below 150 knots. In reference _ the relation- ship of these plotted values to acceptable boundaries is indicated. The comparison, reproduced here in figure 15, indicates that in the approach- speed region the damping is poorer than the acceptable values. It is also noteworthy that with decreasing speed the beneficial effects of the yaw-damper installation tends to diminish until at approach speeds the effect is quite small, the damping being poorer than acceptable with or without the damper. This is consistent with the opinions of the pilots that the dampers were relatively ineffective in this region.

It does not appear that the differences in damping between the basic airplane and the airplane with tanks on are large enough to have affected the approach speed greatly. In particular, the fact that the damping was better with tanks on than off at 12_ knots (damper off) combined with the fact that the pilots did not discern an improvement with the damper on argues that damping could not have been a primary factor in influencing approach speed.

l0

CONCLUS 10NS

Flying-qualities studies were conducted in flight on a tailless

delta-wing fighter-type airplane in the landing-approach configuration

in order to investigate in more detail the factors that contributed to

the pilots' selection of a landing-approach speed. The following factors

were found to be significant:

i. In abrupt aileron rolls the airplane tended to roll around the

highly inclined longitudinal axis so that significant sideslip angles

developed in the roll as a result only of this kinematic effect. This

would augment the usual adverse yaw charactecistics of the ailerons which

were powerful enough that the airplane would yaw and not roll in response

to slow aileron movements.

2. The rolling response to the ailerons was greatly reduced as a

result of the dihedral effect operating at the kinematically developed

sideslip angles.

3. With tanks installed the landing-approach speed was higher than

it was for the basic airplane. With tanks, _udder-free directional

instability (i.e., a rudder-lock) occurred a5 a sideslip angle of 6 ° as

the airspeed was reduced from 135 to 125 kno_s, a condition that was not experienced on the basic airplane. This sideslip angle of 6 ° could be generated in moderate aileron rolls as a res_xlt of kinematic effects mentioned above.

4. The special significance of these f;_ctors lies in the fact that collectively they can occupy so much of the i?ilot's attention that he has reduced capability of coping with the problems of precise flight-path control and, accordingly, he demands a great._r speed margin above the stall to allow for airspeed fluctuations.

Ames Research Center National Aeronautics and Space Administration Moffett Field, Calif., Jan. 15, 1939

ll

REFERENCES

1. White, Maurice D., Schlaff, Bernard A., and Drinkwater, Fred J., III:

A Comparison of Flight-Measured Carrier-Approach SpeedsWith Values

Predicted by Several Different Criteria for 41 Fighter-Type Airplane

Configurations. NACA RMA57Lll, 1958.

o Miller, J. C.: Control System Calibration of F4D-1 Airplane, BUAER No. 130740. Rep. CG-MR-160_ Douglas Aircraft Co., Inc., E1 Segundo Division, Aug. 19_.

o Drinkwater, Fred J., III, and Cooper, George E.: A Flight Evaluation of the Factors Which Influence the Selection of Landing Approach Speeds. NASA MEMO I0-6-58A, 1958.

e McNeill, Walter E., and Vomaske, Richard F.: A Flight Investigation to Determine the Lateral 0scillatoryDamping Acceptable for an Airplane in the Landing Approach. NASA MEMO 12-I0-58A, 1959.

TABLEI.- TESTAIRPLANE CHARACTERISTICS

Engine

Type • • • • @ • J57-P8-A Maximum th_st'wlthout'_te_u_er'(noA_); lb ....

10,200 Maximum thrust with afterburner (nominal), ib ......

16,000 Fuel regulator ..................... JFC 12-2 Airplane Fuselage Length, ft ......................

38.63 Wing Airfoil section Root ....................

NACA 0007-63/30 -9030 ' Modified Tip ....................

NACA 0004.5-63/30 -9030 ' Modified Span, ft .......................

33.5 Area, sq ft ......................

557.0 Taper ratio ......................

0.332 Aspect ratio .....................

2.02 Mean aerodynamic chord, ft ..............

18.25 Leading-edge sweep, deg ................

52.5 Elevon Area, sq ft (total) ..................

45.14 Pitch trimmer Area, sq ft (total) .................. 26.84 Vertical tail Area, sq ft ......................

47.7 Span, ft .......................

7.58 Rudder Manual Area, sq ft .....................

i0.7 Yaw damper Area, sq ft .....................

5.5 Gross weight as tested without external w_ng tanks Empty ...................

15,870 dingil;O; ibf el) ................

16,870 Gross weight as tested with external wing tanks Landing (lO001b fuel) ................

17,260 z3 ,-I !

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