NASA Technical Memorandum 81972 lNASA-TM-81972 19820005275

_or z,o_. r_t_v _e.o_ _s ROo_Limited Evaluation of an

F-14A Airplane Utilizing anAileron-Rudder Interconnect

Control System in the

i Landing Configuration _g_'_

Wendell W. Kelley and Einar K. Enevoldson _'- _ _ _'_

., _._-_ _ _,_, ._,_,G_._"_ _q, _ _.,_\l_

x,r,' _\?,,q_'"... q\_.-_"

DECEMBER 1981

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https://ntrs.nasa.gov/search.jsp?R=19820005275 2019-12-30T23:39:18+00:00Z

NASA Technical Memorandum 81972

Limited Evaluation of an

F-14A Airplane Utilizing anAileron-Rudder Interconnect

Control System in the

Landing Configuration

Wendell W. KelleyLangley Research CenterHampton, Virginia

Einar K. Enevoldson

Dryden Flight Research CenterEdwards, California

NILSANational Aeronauticsand Space Administration

Scientific and Technical

Information Branch

1981

SUMMARY

A flight test was conducted for preliminary evaluation of an aileron-rudder

interconnect (ARI) control system for the F-14A airplane in the landing configura-

tion. Two ARI configurations were tested in addition to the standard F-14 flight

control system. Flight data and pilot comments were used to compare the effects of

each control system on airplane lateral-directional handling qualities during a

final-approach-course correction maneuver.

Results of the flight test showed marked improvement in handling qualities when

the ARI systems were used. Sideslip due to adverse yaw was considerably reduced, and

airplane turn rate was more responsive to pilot lateral-control inputs. Pilot com-

ments substantiated the flight data and indicated that the ARI systems were superior

to the standard control system in terms of pilot capability to make lateral offset

corrections and heading changes on final approach.

INTRODUCTION

In cooperation with the Department of the Navy, the Langley Research Center and

the Dryden Flight Research Center have conducted simulation and flight tests to

investigate improvements to the handling characteristics of the F-14A airplane in the

landing configuration. The study was undertaken as a result of fleet-pilot comments

and quantitative flight data which indicated some undesirable lateral-directional

handling qualities in the landing configuration. Specifically noted were the gener-

ation of a significant amount of adverse yaw following lateral stick inputs and a

lightly damped Dutch roll mode, both of which made it difficult to control airplane

heading precisely in a high-workload task such as a carrier landing.

A previous simulation study, described in reference I, investigated the use of

an aileron-rudder interconnect (ARI) system which involved extensive modifications to

the roll and yaw control systems of the fleet airplane. Two ARI configurations were

tested, and results were compared with those of the standard control system. Side-

slip due to adverse yaw was considerably reduced by the ARI systems, and pilots were

able to control heading more precisely.

The flight test described in this report was conducted to determine whether

similar improvements in handling characteristics could be demonstrated during actual

flight tests using the ARI systems. The test was conducted at the Dryden Flight

Research Center by using an F-14A airplane modified with the experimental control

systems.

SYMBOLS

All aerodynamic data and flight motions are referenced to the body system of

axes shown in figure I. The units for physical quantities used herein are presented

both in the International System of Units (SI) and in U.S. Customary Units. The

measurements and calculations were made in U.S. Customary Units. Conversion factors

for the two systems are given in reference 2.

ay lateral acceleration, positive along positive Y body axis, g units

g acceleration due to gravity (lg = 9.8 m/sec2), m/sec2 (ft/sec2)

Ix,Iy,IZ moments of inertia about X, Y, and Z body axes, respectively,kg-m2 (slug-ft2)

IXZ product of inertia with respect to X and Z body axes, kg-m2 (slug-ft2)

Mi indicated Mach number

p airplane roll rate about X body axis, deg/sec or rad/sec

q airplane pitch rate about Y body axis, deg/sec or rad/sec

r yaw rate about Z body axis, deg/sec or rad/sec

s Laplace variable, 1/sec

u,v,w components of airplane velocity along X, Y, and Z body axes,respectively, m/sec (ft/sec)

V airplane resultant velocity, m/sec (ft/sec)

Vi indicated airspeed, knots

X,Y,Z airplane body axes (see fig. I)

angle of attack, deg

_ias signal used to bias _. in ARI control systems degl

e effective angle of attack, _.z+ _ias' deg

_. indicated angle of attack, degl

angle of sideslip, deg

6 differential horizontal-tail deflection, positive foraright roll, deg

6h symmetric horizontal-tail deflection, positive for airplanenose-down control, deg

6h,DLC incremental horizontal-tail deflection due to DLC control inputs,positive for airplane nose-down control, deg

6h,p symmetric horizontal-tail deflection commanded by pilot longitudinalstick deflection, positive for airplane nose-down control, deg

6ped rudder-pedal deflection, positive for right yaw, cm (in.)

6r rudder deflection, positive for left yaw, deg

6 pilot longitudinal stick deflection, positive for pitch-up, cm (in.)s,p

2

6 pilot lateral stick deflection, positive for ri_t roll, cm (in.)sir

6 symmetric spoiler deflection (both wings) due to DLC control inputs,sp,DLC positive for upward deflection, deg

6 left-wing spoiler deflection due to lateral stick inputs, positive forsp,L upward deflection, deg

6 ri_t-wing spoiler deflection due to lateral stick inputs, positive forsp,R

upward deflection, deg

()6sp'LT total left-wing spoiler deflection, positive for upward deflection, deg

()6sp'RT total ri_t-wing spoiler deflection, positive forupwarddeflection,deg

6_ DLC th_bwheel deflection, positive for aft deflection, deg

Euler angle, deg

_breviations:

AGL above ground level

ARI aileron-rudder interconnect

DLC direct lift control

LSRI lateral-stick-to-rudder interconnect

SAS stability augmentation system

DESCRIPTION OF AIRPLANE

The F-14A is a two-place, twin-engine, jet fighter airplane having a variable-sweep wing and twin vertical tails. A photograph of the test airplane is shown infigure 2. In the landing configuration wing sweep is fixed at 20°, whereas in themaneuvering configuration sweep varies from 20° for low subsonic speeds to 68° forhigher speed flight. Each wing is configured with leading-edge slats, trailing-edgeflaps, and four upper surface spoiler panels. Figure 3 shows details of the wingflap, slat, and spoiler arrangements. In the landing configuration, leading-edgeslats are deflected 17° and trailing-edge flaps are deflected full down to the 35°position; whereas the spoilers are normally raised to the 3° position. A speed brakeon the upper and lower surfaces of the aft fuselage provides increased drag andallows the use of higher engine thrust settings for the landing approach.

Empty weight of the airplane used in the flight tests was 198 084 N (44 531 ib).Fuel weight during the tests varied from 32 027 N (7200 ib) to 13 345 N (3000 Ib).Table I lists the mass and dimensional characteristics of the airplane at a fuelweight of 17 793 N (4000 ib).

BASIC AIRPLANE FLIGHT CHARACTERISTICS

Fleet-Pilot Comments

The following summary of F-14 lateral-directional handling characteristics wasobtained from interviews with several F-14 qualified Navy pilots, some with test-pilot experience. The pilots were asked to comment on airplane handling qualitiesduring all phases of the approach and landing, particularly during day and nightcarrier approaches.

According to those interviewed, the handling-qualities problems in the landingconfiguration are primarily due to adverse yaw following lateral stick inputs and alightly damped Dutch roll mode. These effects are apparent both in visual andinstrument tasks. Unless the pilot applies a generous amount of coordinating rudder,there is substantial adverse yaw both rolling into and out of turns. The adverse yawand resultant heading excursions cause difficulty both in holding a specific headingand in making precise lineup corrections during an approach. The result is thatpilot work load, already heavy because of the difficulty of a carrier approach,increases considerably because of the adverse control and damping characteristics.

Although the problems exist in all phases of the approach, the primary area ofconcern is from an altitude of approximately 61 m (200 ft) down to touchdown, sincesmall lineup errors become important and pilot gain goes up considerably in the ter-minal phase of the approach and touchdown. Turbulence adds a further level of diffi-culty to the problems.

Flight Data

Analysis of flight data in the landing configuration tends to reinforce thequalitative pilot comments. Figure 4 shows the response of the F-14 test airplane toa lateral stick input applied to perform a bank-angle reversal from 30° to -50°.Rudders were not used. Note that a step lateral stick input of approximately 1.9 cm(0.75 in.) resulted in a 12° sideslip angle <8> which reached peak amplitude 3.0 secafter the lateral-control input. Note also the 1.5-sec delay in yaw-rate response_r_ before the airplane began to turn in the proper direction; furthermore, the ini-tial response appears to be opposite to the direction of stick input. Both of thesecharacteristics contribute greatly to difficulties in precise heading control andtend to substantiate the pilot comments.

To eliminate effects of the stability augmentation system (SAS), both the rolland yaw SAS were disengaged for the test shown in figure 4. As a result, the Dutchroll motions are more lightly damped than motions the airplane would exhibit if theSAS system were engaged. However, the adverse yaw characteristics of the F-14 areessentially identical whether the SAS is engaged or disengaged.

DESCRIPTION OF FLIGHT CONTROL SYSTEMS

Basic System

The basic flight control system of the flight-test airplane, designated controlsystem A in this report, was identical to fleet F-14 systems. The system consistedof mechanical linkages, spring and bobweight feel devices, hydraulic actuators, and astability augmentation system (SAS). The SAS functions were generated by roll,

4

pitch, and yaw computers which responded to inputs from the pilot's controls andvarious stabilization sensors. The computer outputs were fed through SAS actuatorswhich drove the control-system mechanical linkages to produce surface motions. TheSAS inputs were in series with pilot inputs and did not produce control-stick motion.

Pitch control.- Figure 5 shows a schematic diagram of the longitudinal (pitch)channel of the airplane. Pilot inputs at the stick provided a direct mechanicalinput to the all-movable horizontal tail (stabilizer). A washed-out pitch-rate feed-back signal provided stability augmentation. The two inputs combined to produce

symmetric stabilizer deflections _(6_ over a range from 10° to -33°. The pitch SAShhad an authority limit of ±3° deflection, and deflection rates were limited to20 deg/sec.

Direct-lift control.- For improved flight-path-angle response, the pilot couldselect direct lift control (DLC) with the control-stick DLC switch. A thumbwheel,

also located on the control stick, provided for DLC inputs and was spring loaded tothe neutral position. When DLC was engaged, all spoilers extended 3° above thecruise zero-deflection position. (See fig. 3.) Forward rotation of the thumbwheelextended the spoilers and aft rotation retracted them according to the schedule shownin figure 6.

The trailing edges of the horizontal stabilizers were displaced 6° downward fromtheir trim position when DLC was engaged in order to compensate for the change inpitching moment due to spoiler extension. (See fig. 6.) When the thumbwheel controlwas rotated fully forward, the spoilers extended to the 12° position and the stabi-lizer trailing edges were displaced 8° down. This action increased the rate ofdescent. When the thumbwheel control was rotated fully aft, the spoilers retractedto the -4.5° position and the stabilizer trailing edges returned to the trim posi-tion. This action decreased the rate of descent.

If DLC was not selected, all the spoilers were deflected 4.5° downward fromtheir cruise zero-deflection position when trailing-edge flaps extended beyond 25°.(See fig. 3.) In this mode of operation, spoilers were used only for roll control.

Lateral control.- The lateral-control system used a combination of differentialhorizontal-stabilizer deflection and spoiler deflection for roll control. Figure 7

shows a schematic diagram of the roll channel. Without roll SAS engaged, 6adeflection was controlled solely by mechanical inputs from the control stick and hada maximum deflection of ±7°• When roll SAS was engaged another ±5° deflection was

provided through a series actuator, which received electrical inputs from a lateral-control stick-deflection sensor and stabilization signals from the roll gyro.

Whether roll SAS was engaged or disengaged, spoilers were used to assist rollcontrol and were actuated via electrical signals from a lateral stick-deflectionsensor. Figure 8 shows a schematic diagram of the spoiler control system, includingboth roll and DLC inputs. Lateral-stick-to-spoiler gearing schedules, which were afunction of whether DLC was engaged or disengaged, are shown in figure 9.

Directional control.- Conventional rudders on each vertical tail were used for

directional control. Figure 10 shows a schematic diagram of the yaw channel of theairplane. Full rudder authority (±30°) was available to the pilot through a mechan-ical linkage to the rudder pedals. With yaw SAS engaged, stabilization signals froma yaw rate gyro and lateral accelerometer were blended with pilot rudder inputs.

5

Experimental Systems

The ARI control systems which were evaluated in the current flight test resultedfrom modifications to a high-angle-of-attack (_) ARI design which is described inreferences 3 and 4. The high-_ system was developed by the Langley Research Centerto improve stability and control characteristics of the F-14 during high-_ maneu-vering. A subsequent simulator study (ref. I) examined the feasibility of applyingthese ARI concepts to the landing configuration, and at the conclusion of that studytwo ARI configurations were selected for flight evaluation. Designated controlsystem B and control system C, the ARI configurations represented two differentlevels of modification to the roll and yaw channels of the high-e ARI design. Thepitch channel, however, remained unchanged and was identical in all three controlsystems flown (A, B, and C).

Control system B.- The ARI system, shown schematically in figures 11 and 12,featured several modifications to the roll and yaw channels, respectively, of controlsystem A. The primary modifications to the roll channel were as follows: (I) aprovision to fade out differential tail deflection due to lateral stick inputs asangle of attack increased, and (2) increased roll-rate damping. The yaw channel alsofeatured the following two modifications: (I) a lateral-stick-to-rudder (LSRI)interconnect gain to counteract adverse yaw automatically, and (2) a roll-rate feed-back to increase lateral-directional damping.

The angle-of-attack and Mach scheduling features of the ARI were designed toimprove the handling qualities of the airplane during high-_ maneuvering and toreduce the probability of control-induced spin entries. However, at approach andlanding conditions (_. = 12°), these ARI features were ineffective. In order to

• lutilize the ARI durlng landing it was necessary to bias the airplane indicated angle-

of-attack signal (_i), which was used as an input to the control systems. To accom-plish this, a positive angle-of-attack bias signal (_bias = 8"96°) was summed with_. to produce larger effective angle-of-attack values (_e) that were then used as• l

inputs to the e schedules. As a result, ARI features became effective at all air-plane angles of attack above approximately 3° and, therefore, were active during thelanding approach. The relationship between true angle of attack _ and the angle of

attack sensed by the test airplane _ vane (_i) is shown in figure 13. Note thatapproach and landing operations always occurred at Mi < 0.55 and, thus, the Machscheduling feature did not vary control-system gains in the landing configuration.

Control system C.- For control system C the same value of _bias was used as incontrol system B; however, several changes were made to the fixed gain values.Figure 14 shows the yaw channel of control system C. The only gain change from con-trol system B was an increase in the LSRI gain from 2.60 to 3.19. This providedslightly more rudder for turn coordination and resulted in quicker yaw response forinitiation of turns.

The roll channel of control system C is shown in figure 15. Two changes weremade from control system B. First, the stick-to-differential-tail gain was slightlyincreased to match the increased LSRI gain. This was accomplished by moving the

breakpoints in the 6a/6srr_ loop from 14" and 40" to 17° and 43°, respectively. Theother gain change was an increase in the roll-rate damping multiplier from 4.0 to

5.0, which provided better damping to complement the increased LSRI and 6a161gains, s,r

6

TEST PROCEDURES

The flight-test airplane was an F-14A which was modified to include the ARIsystems described previously. By using cockpit switches, the pilot could select thebasic F-14 control system (A) or either of the two ARI configurations (B or C). Theairplane was flown by a NASA research test pilot, and the tests were conducted at theDryden Flight Research Center, Edwards, California.

Since it was not possible to conduct actual carrier landings at the test site,it was difficult to achieve the level of task difficulty which would allow direct

comparison of flight results with those of the simulation evaluation described inreference I. It is generally recognized that carrier landings, especially at night,present a higher level of difficulty than runway landings since carrier approachesare flown to a touchdown point which is moving relative to the approaching airplane.Lateral lineup cues are quite different from those present in approaches to a runway,and the techniques used to accomplish the lineups are generally more demanding.

The flight-evaluation task was a series of 12 approaches flown to the primaryrunway at Edwards Air Force Base, California. In order to increase the task diffi-culty, the approaches were accomplished from a lateral offset on final approach.Landing gear, landing flaps, and speed brakes were extended for all approaches; DLCwas also selected. From an altitude of 200 m (656 ft) AGL on downwind, the pilot

began a 180° left descending turn to final approach while slowing to approach angleof attack. The turn to final approach was completed at approximately 100 m (328 ft)AGL and 2000 m (6562 ft) from the runway threshold. Laterally, the turn was com-pleted by rolling out 25 m (82 ft) to the left of the runway center line, lining upon a painted stripe on the runway. This offset was maintained down to an altitude of50 m (164 ft) AGL, at which point a lateral correction to center line was initiated.The lineup correction had to be made rapidly enough to enable landing on the runwaycenter line at the nominal touchdown point (approximately 300 m (984 ft) from thethreshold). Each approach was terminated at an altitude of 20 m (66 ft) AGL and notouchdowns were made. All approaches were flown in day visual-flight-rule (VFR)conditions.

The purpose of using the offset approaches was to require a maneuver which wouldidentify lateral-directional handling deficiencies. The lateral correction to centerline required the pilot to make a fairly rapid series of lateral stick inputs.Rudder-pedal inputs were used only when necessary to complete the maneuver.

A total of 12 offset approaches were flown, with 4 approaches in each control-system configuration. The sequence began with four approaches using control systemB, followed by four approaches using control system A, and finally with fourapproaches using control system C.

Fuel weights during the first and last approach were 32 027 N and 13 345 N(7200 ib and 3000 Ib), respectively. Since this involved only fuselage fuel tanks,changes in moments of inertia and center of gravity were negligible.

DISCUSSION OF RESULTS

Time histories are shown in figures 16, 17, and 18 for an offset approach withcontrol systems A, B, C, respectively. Only one approach is shown for each controlsystem. The three approaches selected were typical in terms of the results noted foreach particular system.

7

Control System A

Figure 16 shows the time history of a typical approach with control system A.The time scale shown at the bottom of the figure represents elapsed time from anarbitrarily selected point prior to the offset correction maneuver. The point atwhich the pilot applied a lateral stick input to begin the correction to center lineoccurred at approximately 16 sec on the time scale. At that time the pilot applied aright lateral stick input of approximately 2.5 cm (1.0 in.) for 0.5 sac. This wasfollowed by several control-stick inputs to stabilize the maneuver and to roll outwith wings level on the runway center line. The maneuver lasted 12 sec, measuredfrom the time the offset correction was initiated until the minimum altitude occurred

which resulted in termination of the approach.

Analysis of the motions shows a noticeable lack of yaw-rate response for almost3 sac after the initial control input, indicating adverse yaw. In addition, signifi-cant sideslip (8) excursions occurred throughout the maneuver which is indicative ofboth adverse yaw and a lightly damped Dutch roll mode. Pilot roll-control inputs, aswell as roll and sideslip oscillations, were of significant magnitude throughout themaneuver.

Control System B

Figure 17 shows time-history results of an approach with control system B. Theinitial lateral input was very similar to that used on the previously describedapproach; however, the magnitude of subsequent inputs tended to decrease as themaneuver progressed. Following the application of lateral-control inputs, there wasalmost immediate yaw-rate response in the proper direction, and yaw oscillations werewell damped throughout the maneuver. Very little sideslip occurred as a result ofthe control inputs. Since pilot rudder inputs were not used, these results indicatea favorable effect of the lateral-stick-to-rudder feature of the ARI design. Duringthe 12-sec maneuver it is apparent that pilot roll-control inputs, as well as rollrate and sideslip oscillations, were smaller in magnitude and more damped than thoseof control system A.

Control System C

Figure 18 shows time-history results of an approach with control system C. Theinitial lateral-control input was smaller than in the previous two cases, but subse-quent inputs were very similar to the case for control system B.

Airplane yaw-rate characteristics were approximately the same as that noted withsystem B in terms of response to controls and damping. However, the magnitude of yawrate relative to size of control input was slightly higher than for control system B,because of the higher lateral-stick-to-rudder gain in control system C. The effectof this gain on rudder response is shown in the time-history traces for both controlsystems B and C. Sideslip-angle oscillations were very small throughout the maneuverand similar to the results noted for control system B. Also, pilot inputs and rollrate were well damped as with control system B.

8

Pilot Comments

The following statements are a summary of pilot comments which were made at aflight debriefing immediately following the completion of the test flight.

Control system A.- There was a distinct, dramatic difference in the airplaneresponse between the basic control system (system A) and the ARI systems. The noseyawed back and forth excessively following lateral-control inputs with controlsystem A. It was not possible to do the offset maneuver with feet on the floorbecause of the adverse yaw and lightly damped Dutch roll motions. After trying thefirst maneuver without using rudder pedals, the subsequent maneuvers were done withrudder-pedal assistance. There was significant adverse coupling between roll andyaw, making it difficult to control heading with lateral stick inputs. Thus, it wasnecessary to coordinate a large amount of rudder-pedal input to keep the nose turningin the proper direction.

Control system B.- The airplane response to controls was good and the lineuptask was much easier with this system. It was not necessary to use rudder pedals forany part of the maneuver. The initial roll response was a little jerky, but this wastrue of each system tested including control system A.

Control system C.- Response to control inputs was very good and there were noobjectionable qualities. Again, the use of rudder pedals was not necessary for theARI system. Although the pilot did not encounter a lateral PIO (pilot-inducedoscillation) with this system, there might possibly be a PIO problem with a realhigh-gain pilot because of the rapid roll and yaw response. Control system C did notseem dramatically better than control system B, but there was a tremendous improve-ment over control system A.

CONCLUDING REMARKS

A previous carrier-landing simulation study of the F-14A airplane concluded thatan aileron-rudder interconnect (ARI) control system produced better handling quali-ties and improved pilot performance over the standard F-14A control system. Theobjective of the present flight-test program was to determine whether similarimprovements in handling qualities could be demonstrated in a real airplane config-ured with the same control systems.

Although the approach task was different in the simulation and flight test,strong similarities were exhibited in the results and pilot comments. Flight-testresults of the unmodified F-14A showed considerable adverse yaw, large sideslip

excursions, and lightly damped Dutch roll oscillations. On the other hand, both ARIconfigurations tested provided a marked improvement in handling qualities as follows:airplane turn rate was more responsive to pilot lateral-control inputs, sideslip wasreduced because of the coordinating rudder feature of the ARI, and the Dutch rolldamping was significantly increased.

It is emphasized that the flight test conducted was of a very limited nature, sothat no final conclusions can be drawn regarding the suitability of the ARI systemsfor actual carrier landings. However, the results showed that the ARI systems pro-vided very definite improvements in handling qualities in areas where the basic

airplane is known to be deficient. In order to investigate fully the improvementspossible with the ARI systems, further flight tests should be conducted to optimizethe systems and evaluate them during actual carrier landings.

Langley Research Center

National Aeronautics and Space AdministrationHampton, VA 23665November 9, 1981

REFERENCES

I. Kelley, Wendell W.; and Brown, Philip W.: Simulator Results of an F-14A AirplaneUtilizing an Aileron-Rudder Interconnect During Carrier Approaches and Landings.NASA TM-81833, 1980.

2. Mechtly, E. A.: The International System of Units - Physical Constants and Con-version Factors (Second Revision). NASA SP-7012, 1973.

3. Nguyen, Luat T.; Ogburn, Marilyn E.; Pollock, Kenneth S.; Deal, Perry L;Brown, Philip W.; and Whipple, Raymond D.: Piloted Simulator Study of High-Angle-of-Attack Characteristics of F-14 Airplane With Maneuver Slats. NASATM-80058, 1979.

4. Nguyen, Luat T.; Gilbert, William P.; Gera, Joseph; Iliff, Kenneth W.; andEnevoldson, Einar K.: Application of High-_ Control System Concepts to aVariable-Sweep Fighter Airplane. AIAA-80-1582, Aug. 1980.

10

TABLE I.- MASS AND DIMENSIONAL CHARACTERISTICS OF TEST AIRPLANE

Weight with 17 793 N (4000 Ib) of fuel, N (ib) .................... 215 877 (48 531)

Moments of inertia, kg-m2 (slug-ft2):

Ix ............................................................... 89 647 (66 120)Iy ............................................................. 360 217 (265 681)Iz ............................................................. 444 288 (327 689)IXZ ................................................................ -3440 (-2537)

Wing dimensions:Span, m (ft) ................................Area, m2 (ft2) [ [ ii ii[ii[[[[i [ [ i [[[ [ [ [ 19.55 (64.13).................. 52.5 (565)Mean aerodynamic chord, m (ft) ....................................... 2.99 (9.80)

Center of gravity, percent of mean aerodynamic chord .......................... 13.4

11

qY

v

w\\\\\ u\

P

r _ 6 \X\ X

V

Figure I.- The body system of axes.

12

L-81-236

Figure 2.- Photograph of the F-14A flight-test airplane in thelanding configuration.w

Spoi lers

11--12 ° (DLC engaged, thumbwheel full forward)

Leading-edge "_ _--3 o (DLC engaged, thumbwheel neutral)slats down

(17o) -4.5 o (DLC engaged, thumbwheel full aft;or DLCnot engaged)

Trailing-edgeflaps down

(35o)

Figure 3.- Wing control surfaces.

5oE"

t --5o i I I

40

P,deg/sec 0

-40

6a, deg 0

-4 I I I I I I I

_s,r,cm05_-5 i I I I I I I I

B, deg 0

-2o I I I I I I I I I

r,

deg/sec O_-lo I I0 2 4 6 8 10 12 14 16 18

Time, sec

Figure 4.- Responseof the F-14 test airplaneto lateralstick input.

15

d DLC DLC engaged

horizontal6TW stab. gearing _'-_'-

_ (See fig. 6)

Power actuator

"30F_-20

+ 1 Deflection

X-10h 6h'P +_ _ O.05s + l -- limit, -L_ 6has,P __ Ol + i0°, _330

101 Rate limit, 36 deg/sec-10-5 0 5 10 15

cms,p' SAS actuator

Deflectionlimit, Rate limit,

+20 deg/sec+3°

Limit, 2s Limit, I

q _ _+50 _ +10° d O.3 ,deg/sec 2s + i -

Figure 5.- Schematic diagram of pitch channel of all control configurations.

15 -

IO-

6sp,DLC' 5 --deg

0-

-5 _

-10

-75 -50 -25 0 25 50

_STW,deg

(a) Spoiler deflection as a functionof DLC thumbwheel inputs.

10 --

8

6

6h,DLC'

deg

4

2

0 l-50 -25 0 25 50

Forward Aft

6TW, deg

(b) Horizontal-stabilizer deflection as a

function of DLC thumbwheel inputs.

Figure 6.- Control-surface deflections for DLC operation.

17

To spoiler control Power actuator

I I i LDeflectionLimit, 0.7874 + 0.05s + 1 r limit,_s,r "-- _+8.89 cm _ _" _ I _

+ Rate limit, 36 deg/sec +12o - a

. SAS actuatorI

JI

1.685 Limit, + _._ Rate limit, Deflection--_ _ -_ limit,

0.5s + 1 +10o +33.4 deg/sec +5o

, T I J

Limit' _-0"04, IP' < 135 deg/sec_ Limit, _

p _ +250deg/sec -0.15, IP] > 135 deg/sec +15.2 °

Figure 7.- Schematic diagram of roll channel of control system A.

Right-wingspoiler actuators

DLC DLC _ 1 1 H Deflection I_ (G)

spoiler engaged 6sp,DLC + O.05s + I limit6TW _ gearing ' sp,R

DLC (see fig. 6) _ + 150Ratedeg/seclimit' 55o,_4.5 o Tengage T J

i Left-wi ng

switch T 6sp,R 6sp,DLC spoil er actuators

_ Lateral + I I __ (G)

spoiler O.05s + I Deflection6s,r gearing + Rate limit, limit, sp,L

(see fig. 9) 6sp,L 150 deg/sec 55o,-4.5 o T

Figure 8.- Schematic diagram of spoiler control system of allcontrol configurations.

60--

I I I I-10 -5 0 5 10

6s cm_r _

(a) DLC engaged.

Figure 9.- Spoiler deflectionsas a functionof lateralstick input.

2O

60--

sp,L 30 --or

sp,R' 6sp 6sp,Rdeg

20--

I I I I-10 -5 0 5 10

6 cms_r _

(b) DLC not engaged.

Figure 9.- Concluded.

21

Power actuator

, O.05s + 1 DeflectionLimit + _ = limit, _ .__6 r

6ped -3.94 _ -+30o + Rate limit, +_30o106 deg/sec

s m+_50 SAS actuator2s + 1 deg/sec

I Rate _ Deflecti°n

+_ _ limit, limit,+ 63.4 deg/sec +_19°

s<cmay ' ,r 1 Limit,22.5, las,rl >_2.54 cm o.05s + i +_21.5o

Figure 10.- Schematic diagram of yaw channel of control system A.

Power actuator

_ To spoilercontrol + _ _ _

Deflection+_t< 1

Limit, 0.7874 _ m- limit, _ _a_s,r *_8.89cm O.05s + i +_120

Rate limit,36 deg/sec

1! 18 , 40

Limit, I" I_'/0.7874 +_70 .~ _/

7 14

_e ' deg

SAS actuator

. 7 14 "_.

1.685 Limit, + Rate Iimit, Deflectionlimit,

O.5s + I +-10° 0 18 40 +33.4 deg/sec _+50

_e ' deg

°°4P<1311L eOOdeg/sec .,<+

4 -- ---o-0.15, lpl•135 + -

0 OIdeg/sec .65 6.35 _e' de9 39

Mi [6pedl , cm

Mi I_pedI--'t °e

Deadband, .----o_

r _ +-200

deg/sec _ • 44oe -

Figure 11 .- Schematic diagram of roll channel of control system B.

&o

Power actuator

eeconLimit, _ -3.94 O.05s + 1 limit, 6

6ped -+7.62 cm m _ rRate limit, -+30°106 deg/sec

20 22 SAS actuator i

,6s r 2.60 _, 1 i_. / _/ _ limit, _ limit,

' _' 0.125s + 1 i_/_/ 63.4 deg/sec _+19oi 0i0 12

_e ' deg

P -+30 deg/sec

_e < 44o

Mi < 0.55

r ._ 2s Limit, I2s + I _+50o

-F

_ 9.15, 16 I < 2.54 cm

ay s ,r I Limit,

22.5, I_s,rI >__2.54 cm o.05s + i +21.5 o

Figure 12.- Schematic diagram of yaw channel of control system B.

24

20

16

c_, deg

12

8

4

i0 4 8 12 16 20 24 28

_i, deg

Figure 13.- Relationshi 9 of true angle of attack to indicated

angle of attack.

25

o_

Power actuator

i I Iki mit, + IDefl e cti on

_e_--__-_.0_-_-_ __-_ 0.0_,,[_m_,___cm Rate limit, -*30° r

106 deg/seci

20 22 SAS actuator

_k _/_/6s,r_ 3.19 I_7¢7 Rate I IDeflectionllmlt, H limit,

o. 125s + I I _'_'/_'/ 63.4 deg/sec H +19°0 I0 12

me , deg

Limit, _me! C_eJ LM

P _ +-30 I

deg/sec 44o

Mi < 0.55

2s Limi t,r_2s + 1 +bu°-_

+I

+

9.15,16s,rl < 2.54 cm 1 Limit,ay

22.5,16s,rl > 2.54 cm O.05s + 1 +21.5°

Figure 14.- Schematic diagram of yaw channel of control system C.

J, To spoiler control Power actuator

[ il ILimit, = 0.7874 + =+__s,r = +8.89 cm _- 0.05s + 1 limit, _ _a

Rate limit,36 deg/sec

I L-,., 21 43

I:A_/Limit,

10 17

_e' deg

_ _ ---Mi SAS actuator

.__ 1_ -,_ Deflection _

,.68_ Limit, I _<"Xb_ _ +33.4Ratedeg/seclimit' limit,O.5s + 1 -+10o +5o

21 43

_e' deg

,, ooI -0.04, IPl 2 135 1 1 "_ %Limit, deg/sec _

q __P _c o_, I_l, 13s _,<_" +39

deg/sec 0 .65 0 _35 O'Mi 1Sped[ , ae, deg

Mi--_ 16pedl_ _e -I

I Deadband, _>_4

r _-- _+200

deg/sec 40e -

Figure 15.- Schematic diagram of roll channel of control system C.

n. cm Begin Terminate5 - 10 maneuver maneuver

5

6s,r 0 - 0

-5

-5 - -I0

25 -

r_

0 __,....._-- ___--__--- __.deg/sec

-25 -

25 -

B, deg 0 -

q_, deg 0 ____- --

-100 I I I I I I I I0 5 10 15 20 25 30 35 40

Time, sec

Figure 16.- Lateral offset correctionmaneuverwith controlsystemA.

28

Begin Terminatemaneuver maneuver

i°E6r' deg -- • _ --_ ----- , '

-30

30

P_

deg/sec -10

-50 I i I I I I I I0 5 10 15 20 25 30 35 40

Time, sec

Figure 16.-Continued,

29

Begin Terminatemaneuver maneuver

180 --

Vi , knots

140 --

I00

2O

_i ' deg

0

90 -

_sp,L) T ,

deg 40 -

-10 --

90 --

(_sp,R) T ,40 --

deg

-10 _- -F- ' l I I -T i I0 5 I0 15 20 25 30 35 40

Figure 16.- Concluded.

3O

Begin Terminatei n. cm maneuver maneuver

6s,r 50 - 0

-5

-5 - -I0

25 --

-25 --

25 --

13, deg

O-- ---- -------- -- ------ _

-I00 I I I I I I I I0 5 10 15 20 25 30 35 40

Time, sec

Figure 17.- Lateraloffset correctionmaneuverwith controlsystemB.

31

Begi n Terminatemaneuver maneuver

-12.5

cm

in. i0 I

1 5

6ped E 0-5

-3 -I0

6r, deg 10 _ _•30

3O

-10deg/sec

-5o I I I I I I I I0 5 I0 15 20 25 30 35 40

Time, sec

Figure 17.-Continued.

32

Begin Termi natemaneuver maneuver

180 -- _

Vi , knots140 -- , _ _

I00 --

20 --

_i' deg 100 --

m

(%p,L)T'deg 4 --

=1 m

9 --

(_sp,R)T '4 --

degI

_'_-- %.-_ l I-1 " I ----I- F-- __ -I , _ , - I0 5 I0 15 20 25 30 35 40

Time, sec

Figure 17.- Concluded.

33

Begin Terminatein. cm maneuver maneuver

(Ss,r0 -- 0

--5 _

-10-5

25 -

r, deg/sec 0 _ _ __ _ .... - ._-- __.

-25

25 --

B, deg

0 _ .... -- _ -_--_-_--_____-- -_

_, dego _- •

-i00 [ ] ] I ] ] I ]0 5 I0 15 20 25 30 35 40

Time, sec

Figure 18.- Lateraloffset correctionmaneuverwith controlsystemC.

34

Begin Terminatemaneuver maneuver

, dega-2.5

I°E6r, deg - -

-30

30 --

P' -10 -deg/sec

-5o I i I I I I I I0 5 I0 15 20 25 30 35 40

Time, sec

Figure 18.- Continued.

35

Begin Terminatemaneuver maneuver

180 -- I I

V., knots1 140 --

, - --_ _

100--

20 --

mi' deg I0 L " v"_V'_- v_'_'_I_'_%/_'IP'_A.__

0 -

90 -

(6sp,L)T '40 --

deg

-10 --

90 --

(6sp'R)T'deg40 --lo - - I I ' - ]- ---_--I--"- - i - -'-i I I0 5 10 15 20 25 30 35 40

Time, sac

Figure 18.- Concluded.

36

1. Report No.

NASA TM-81972I 2. Government Accession No. 3. Recipient's Catalog No.

4. Title and SubtitleLIMITED EVALUATON OF AN F-14A AIRPLANE UTILIZING ANAILERON-RUDDER INTERCONNECT CONTROL SYSTEM IN THELANDING CONFIGURATION

7. Author(s)

Wendell W. Kelley and Einar K. Enevoldson

9. Performing Organization Name and Address

NASA Langley Research CenterHampton, VA 23665

12. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, DC 20546

5. Report Da te

December 1981

6. Performing Organization Code

505-43-13-01

8. Performing Organization Report No.

L-14756

10. Work Unit No.

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Wendell W. Kelley: Langley Research Center, Hampton, VirginiaEinar K. Enevoldson: Dryden Flight Research Center, Edwards, California

16. Abstract

A flight test was conducted for preliminary evaluation of an aileron-rudderinterconnect (ARI) control system for the F-14A airplane in the landing config­uration. Two ARI configurations were tested in addition to the standard F-14 flightcontrol system. Results of the flight test showed marked improvement in handlingqualities when the ARI systems were used. Sideslip due to adverse yaw wasconsiderably reduced, and airplane turn rate was more responsive to pilot lateral­control inputs. Pilot comments substantiated the flight data and indicated that theARI systems were superior to the standard control system in terms of pilot capabilityto make lateral offset corrections and heading changes on final approach.

17. Key Words (Suggested by Author(s))

F-14A airplaneAileron-rudder interconnect (ARI)Flight controlsFlight testFighter airplane

18. Distribution Statement

Unclassified - Unlimited

Subject Category 08

19. Security Oassif. (of this report)

Unclassified

20. Security Classif. (of this page)

Unclassified

21. No. of Pages

37

22. Price

A03

For sale by the National Technical Information Service, Springfield, Virginia 22161

NASA-Lan91ey, 1981

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t Washington,Space20546AdministrationD.C. SpaceNASA-451NationalAdministrationAeronautics andOfficial Business

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