NASA TN D-8026 NASA TECHNICAL NOTE

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MOTION-BASE SIMULATOR STUDY OF CONTROL OF AN EXTERNALLY BLOWN FLAP STOL TRANSPORT AIRCRAFT AFTER FAILURE OF AN OUTBOARD ENGINE DURING LANDING APPROACH

David B. Middleton, George J. Hurt, Jr.,

Hugh P. Bergeron, James M. Patton, Jr.,

Perry L. Deal, and Robert A. Champine

Langley Research Center

Hampton, Va. 23665

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, 0. C. • OCTOBER 1975

https://ntrs.nasa.gov/search.jsp?R=19750024018 2018-05-26T12:22:17+00:00Z

1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.

NASA TN D-8026

4. Title and Subtitle MOTION-BASE SIMULATOR STUDY OF CONTROL 5. Report Date OF AN EXTERNALLY BLOWN FLAP STOL TRANSPORT AIR- October 1975 CRAFT AFTER FAILURE OF AN OUTBOARD ENGINE DURING 6. Performing Organization Code LANDING APPROACH

7. Author(s) 8. Performing Organization Report No.

David B. Middleton, George J. Hurt, Jr., Hugh P. Bergeron, L-9743 James M. Patton, Jr., Perry L. Deal, and Robert A. Champine

10. Work Unit No.

504 -09-31-01 9. Performing Organization Name and Address

NASA Langley Research Center 11. Contract or Grant No.

Hampton, Va. 23665

13. Type of Report and Period Covered

Technical Note 12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration 14. Sponsoring Agency Code

Washington, D.C. 20546

15. Supplementary Notes

16. Abstract

A moving-base simulator investigation of the problems of recovery and landing of a STOL

aircraft after failure of an outboard engine during final approach has been made. The approaches

were made at 75 knots along a 6 0 glide slope. The engine was failed at low altitude and the option

to go around was not allowed. The aircraft was simulated with each of three control systems, and

it had four high-bypass-ratio fan-jet engines exhausting against large triple-slotted wing flaps to

produce additional lift. A virtual-image out-the-window television display of a simulated STOL

airport was operating during part of the investigation. Also, a simple heads-up flight director

display superimposed on the airport landing scene was used by the pilots to make some of the

recoveries following an engine failure.

The results of the study indicated that the variation in visual cues and/or motion cues had

little effect on the outcome of a recovery, but they did have some effect on the pilot's response

and control patterns.

17. Key Words (Suggested by Author(s)) 18. Distribution Statement STOL handling qualities Motion-base simulation Unclassified - Unlimited

Engine-out recovery Motion washout Heads-up flight director Externally blown flap Subject Category 05

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price

Unclassified Unclassified 60 $4.25

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

MOTION-BASE SIMULATOR STUDY OF CONTROL OF AN EXTERNALLY

BLOWN FLAP STOL TRANSPORT AIRCRAFT AFTER FAILURE OF

AN OUTBOARD ENGINE DURING LANDING APPROACH

David B. Middleton, George J. Hurt, Jr., Hugh P. Bergeron, James M. Patton, Jr., Perry L. Deal,

and Robert A. Champine Langley Research Center

SUMMARY

A moving-base simulator investigation of the problems of recovery and landing of a short take-off and landing (STOL) aircraft after failure of an outboard engine during final approach has been made. The approaches were made at 75 knots along a 60 glide slope

and the engine was failed at altitudes of from 38 to 152 in (125 to 500 ft). The pilots were not given the option to go around. The aircraft was simulated with each of three control systems, and it had four high-bypass-ratio fan-jet engines exhausting against large triple-slotted wing flaps to produce additional lift. A virtual-image out-the-window television display of a simulated STOL airport was used during part of the investigation. A simple heads -up flight director display was also superimposed on the airport scene and used during some of the recoveries.

The results of the study indicated that rather sophisticated stability and control sys-tems were required to assist the pilot in maintaining precise control of this airplane after failure of an engine. (Considerable augmentation was also required for normal four-engine landings.) For the "fully augmented" airplane the motion environment affected control very little because very little motion occurred except in turbulence. As the degree of stability and control augmentation was reduced, serious pilot-induced oscilla-tions (PlO's) sometimes resulted from lateral upsets such as strong side gusts or the failure of an engine. These PlO's occurred in both moving-base and fixed-base runs (made for comparison). The initial motion cues following a disturbance felt realistic and assisted the pilot in making rapid and efficient control corrections, but whenever the input-output relationship developed into an oscillatory mode, the PlO tendency became reinforced by lags in the simulation system. In such cases the lags caused the motion to, feel unrealistic and thus degraded the fidelity of the simulation.

When the televised airport scene was added to the simulation, the engine-out (EO) recoveries were often performed more slowly than those that were under strictly instru-

ment conditions. Consequently, maximum excursions from the glide slope and localizer were usually greater, but the recoveries were no less successful. The pilots explained that the gross visual cues from the television (TV) picture reassured them that adequate control of the aircraft was being maintained during the EO transient and recovery and that sufficient time (or distance) remained for unhurried realinement with the runway. When a simple heads-up flight director display was superimposed on the airport scene, the EO recoveries were again performed quickly. The pilots said, however, that in this situation they fixated only on the heads-up director information and tended to ignore the underlying visual scene until the aircraft was under good control and well on its way back to the approach path.

INTRODUCTION

Recent interest in medium-range transport aircraft which can operate out of small airfields has prompted consideration of several aircraft configurations which can produce relatively high lift coefficients at low speeds. For example, during a recent study (ref. 1) the landing speed of a Boeing 707 was reduced approximately 25 percent by use of a powered-lift system in which engine-bleed, compressed air was blown over the upper sur-face of the wing flaps. This condition made it possible to use flap deflections of 700 as compared with 40 0 for plain flaps without boundary-layer control. Handling qualities (particularly those associated with lateral motions) deteriorated noticeably at the lowered speed, and several types of stability augmentation were added or recommended. A sec-ond method of producing high lift at low speeds is with an externally blown flap (EBF) sys-tem where the exhausts from the aircraft's turbofan engines are blown directly against large slotted wing flaps. The failure of one of the outboard engines while in this high-lift configuration, however, will cause significant lift asymmetries as well as thrust vector asymmetries. Consequently, the resulting rolling and yawing moments may create a transient control problem if the handling qualities of the aircraft are not particularly good at low speeds.

The "engine-out" (EO) problems of a conceptual EBF short take-off and landing (STOL) aircraft have been studied analytically, in the wind tunnels (refs. 2 and 3), and in a fixed-base simulation study (ref. 4) at the Langley Research Center. It was deter-mined (ref. 4) that the low-speed handling qualities of the basic airplane would be poor and considerable augmentation would be required in order to obtain a satisfactory pilot rating in both normal and EO situations. However, it was still not known how the presence of cockpit motion might affect the piloting task during EO transients; therefore, the present investigation was undertaken. It was expected that the difficulty of the EO control task would be reduced because the motion environment would provide useful natural cues, and

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that good control could be maintained in normal situations with a much simpler stability and control augmentation system than that used in the fixed-base simulation.

The computer program of the fixed-base study (ref. 4) was adapted to the Langley real-time dynamic simulator (RDS) for a motion study of the EO problem during 60 glide-slope landing approaches at 75 knots. The aerodynamics and control systems were thus essentially the same for both studies and the RDS cockpit controls and instrument displays were arranged like those of the fixed-base cockpit. A virtual-image television scene of a typical STOL airport was used during part of the study. (This same airport scene had been used but presented in a slightly different manner in the study of reference 4. Also, the same pilots that participated in the reference 4 study flew the approaches (using the same techniques) and made the handling-qualities evaluations.

SYMBOLS

In order to facilitate international usage of the data presented, dimensional quanti-ties are presented in both the International System of Units (SI) and U.S. Customary Units. The measurements and calculations were made in U.S. Customary Units. Dots over sym-bols denote differentiation with respect to time.

A,B,C,D,E computer model internal gains

a1 direction-cosine elements of inertial-to-body transformation matrix (1=1,2,3; j=1,2,3)

ay lateral acceleration, g units

at compensated lateral acceleration (ay - YOrOr), g units

b wing span, m (ft)

LCD,ge incremental drag coefficient due to ground effect

LCL,ge incremental lift coefficient due to ground effect

C1 , 1,C12 rolling-moment coefficients

incremental rolling-moment coefficient due to failure of engine 1

Cm pitching-moment coefficient

Cm , o basic pitching-moment coefficient

Cm , 1,Cm, 2 pitching-moment coefficients

Cm,ge incremental pitching-moment coefficient due to ground effect

Cn, 1, Cn, 2 yawing-moment coefficients

L Cn, E1 incremental yawing-moment coefficient due to failure of engine 1

CT thrust coefficient

CX longitudinal -force coefficient

C , o basic longitudinal-force coefficient

CXE1 incremental longitudinal force coefficient due to failure of engine 1

Cx, ge longitudinal -force coefficient due to ground effects

Cy 1 ,C 2 side-force coefficients

zCy,1 incremental side-force coefficient due to failure of engine 1

C Z vertical-force coefficient

basic vertical-force coefficient

C Z ,E1 vertical-force coefficient with only three engines operating

Cz,ge vertical-force coefficient due to ground effects

mean aerodynamic chord, m (ft)

4

forward, middle, and rearward segments, respectively, of the wing flaps

g acceleration due to gravity, rn/sec2 (ft/sec2)

h altitude, m (ft)

Ix'Iy'Tz moments of inertia about XB, YB, and ZB body axes, respectively, kg-rn2

(slug-ft2)

Ixz product of inertia, kg-m 2 (slug-ft2)

K gain

m mass of aircraft

p,q,r rolling, pitching, and yawing angular velocities, respectively, deg/sec or

rad/sec

S wing area, m 2 (ft2)

s Laplace operator

T thrust, N (lbf)

Tc thrust commanded, N (lbf)

u,v,w components of V in the XB, YB, and ZB directions, respectively

uA,vA,wA components of VA in the XB, YB, and ZB directions, respectively

ug,vg,wg random gust velocity along the XB, YB, and ZB body axes, respectively, rn/sec (ft/sec)

9 91 random gust velocity along X1, Y1, and Z1 inertial axes, respectively,

rn/sec (ft/sec)

V ground speed, knots (ft/sec)

VA resultant aerodynamic velocity of airplane, knots (ft/sec)

Vc velocity commanded, knots (ft/sec)

X distance from runway threshold, positive down runway, m (ft)

X',Y',Z' model displacement drive commands to RDS

XB,YB,ZB body coordinate axes fixed at airplane center of gravity

XI, Y1 ,Z 1 inertial coordinate axes fixed at center of runway threshold

Y lateral displacement from localizer, m (ft)

()EO maximum lateral displacement from localizer during recovery after engine-out condition, m (ft)

PVA2SCY = 6r

r 2m

a angle of attack, deg

13 angle of sideslip, deg

6 a aileron deflection, positive for right roll command, deg

Oc control-column deflection, positive for pull force, deg

Of 1 deflection of forward segment of trailing-edge flap, deg

6f2 deflection of second segment of trailing-edge flap, deg

00 total deflection of rearward segment of trailing-edge flap, deg

deflection of from 600, deg (O = -600

6 LT cockpit controller for longitudinal trim

6

OP pedal travel, cm (in.)

6RT cockpit controller for roll trim

rudder deflection, deg

Os asymmetric deflection of spoilers for roll control, positive for right roll command, deg

0s,int asymmetric spoiler deflection beyond which ailerons are slaved to deflect with asymmetric spoiler, deg

Osp symmetrical deflection of spoilers for lift control, deg

horizontal-tail deflection, positive when trailing edge is deflected down, deg

0th throttle deflection, in.

control wheel deflection, deg

(h)Eoaltitude loss with respect to glide-slope beam following an engine-out condi-

tion, m (ft)

localizer error, deg

glide-slope error, deg

zh glide-slope error, m (ft)

0 pitch angle, deg

pitch model angle command or pitch command to RDS

P air density, kg/m 3 (slugs/ft3)

7

T time constant, sec

angle of roll, deg

roll model angle command or roll command to RDS

angle of heading, deg

yaw model angle command to RDS

heading error, deg

[ ] denotes matrix

[ ]i denotes inverse matrix

- aC 1 - acc

- aCy - - -

acz- aCx - - aCm of3

86 6f3 a6'3 O3 ao'3

Cx6sXaC aC aC

a6s- Z6s a6s- m05 - a65

= = aCn 10 ao 5 110s a5 5 'os -

aCxXR

aCz aCm - Z) - m - sp Vsp Jsp au SP 'Isp a6sp

- aCx - aCz - acm X - - mO - aot

- C160

aC1 aC aCy

r 80r 6r aor 0r aor

8

DC1 Clöa = =2.n Cflöa DOa

act C1 =-

aCfl Cfl = PaP!2.

2V 2V

act aCfl

Clr=arb Cfl=

2V

DC OaDOa

aC

2V

DC Yr

2V

DC La Da

acm Cm- - Cm. DCm- ' -iqc (2

2V

Abbreviations:

DLC direct lift control

EO engine out

EBF externally blown flap

IFR instrument flight rules

PlO pilot-induced oscillations

RDS real-time dynamic simulator

SAS stability augmentation systems

STOL short take-off and landing

VFR visual flight rules

9

DESCRIPTION OF SIMULATED AIRCRAFT

The high-wing, T-tail STOL aircraft simulated in this study is essentially the same as the one used in the reference 4 study. A three-view drawing of the aircraft is shown in figure 1 along with assumed full-scale mass and dimensional characteristics. Control-surface deflections and maximum deflection rates are given in table I. Nondimensional aerodynamic characteristics for this aircraft were obtained from the wind-tunnel data curves of references 2 and 3. (See table II.) Also listed in table II are the coefficients (determined from the data curves of ref. 5), representing the so-called "negative ground effects." These ground effects are somewhat less severe than those used in the refer-ence 4 study because they apply to a lower range of CL values.

Four commercial high-bypass-turbofan engines which generate a combined maxi-mum thrust of 147 058 N (33 060 lbf) were simulated. The engine characteristics are shown in table ifi. The engines were mounted to the underside of the wings and canted so that the exhaust impinged directly onto the trailing-edge flap system (see fig. 2) which was deployed for landing. The inboard engines were located at 0.22 semispan and the outboard pair at 0.42 semispan; this arrangement corresponds to the "spread-engine configuration" of references 2 and 4. The "clustered" engine configuration (outboard engines moved in to 0.30 semispan) of these references was also used during a few of the test approaches. This configuration is shown by the dashed lines in figure 1.

The wing incorporated blown (engine-bleed) leading-edge flaps (which were deflected 600) and full-span, triple-slotted trailing-edge flaps (which were set at 6fl/6f2/6f3 = 250/100/600 for the approach and landing condition). A sketch of the flap assembly is shown in figure 2. The first two chordwise elements, f 1 and remained fixed at 25 0 and 10 0, respectively, but the third chordwise element 13 was implemented for active control (operating about the O = 600 reference position). An automatic speed control function was achieved by deflecting all six 13 elements (three on each wing) symmetrically. Either the inboard or the middle f3 elements on each wing were also deflected differentially for aileron roll control. Top-wing spoilers (fig. 2) were geared to the ailerons to produce additional roll control.

Two schemes were used for roll control, one for the clustered-engine configuration and the other for spread engines. The first scheme was to use the spoilers for primary control and gear the inboard 13 elements as ailerons to assist when the spoilers reached ±300. This scheme was used with clustered engines in the reference 4 study and again during the present study. The second scheme (for the spread-engine configuration) was to use the middle 13 elements as ailerons for primary roll control and gear the spoilers to assist whenever the ailerons reached their travel limits of ±200. This scheme is illustrated in figure 3 for a typical case of roll control during a landing approach where

10

the outboard engine on the right wing was failed (at approximately 19 sec). Prior to the engine failure, the ailerons were more than adequate for roll control, but when the failure occurred, they were driven at their maximum rate to their 20 0 limits. (The 6a = _200 in fig. 3 corresponds to left-wing 6f3 = (60 0 - 200) and right-wing of = (600 + 200).) With the ailerons saturated, the left-wing spoilers were deflected an average of about 150 to overcome the remaining unbalanced rolling moment due to the failed engine. Contin-üed roll control by the pilot was accomplished with the spoilers only because the ailerons remained saturated until touchdown.

The rudder on the vertical tail could be deflected ±40 0 for directional control. Leading-edge blowing on the rudder surface was used to increase rudder effectiveness at low speed. (See ref. 2.) The elevators were fixed at 50 0 (trailing edge up) for primary pitch trim of the high-lift landingconfiguration, and the entire horizontal tail (stabilator) was servodriven from the column or pitch-trim controller to achieve pitch control.

STABILITY AND COMMAND AUGMENTATION SYSTEMS

The STOL airplane used in this study was simulated with three different high-gain control systems (selected from ref. 4) having various degrees of stability and command augmentation. These aircraft-control configurations are designated herein as (1) BASIC, (2) BASIC + SAS, and (3) FULLY AUGMENTED. The same throttle gains and engine response characteristics (table ifi) were used for all configurations. Automatic speed control (AUTOSPEED) as depicted in figure 4 was also used with all configurations. The first two configurations involved the airplane with "clustered" engines only whereas the FULLY AUGMENTED airplane was simulated with only the spread-engine locations. The aerodynamic difference of the airplane for these two engine locations is indicated in part (b) of table II. As mentioned previously, the rearward segments of the inboard flaps were used as ailerons on the clustered-engine configuration whereas the rearward segments of the middle spanwise flaps were used as ailerons on the spread-engine configuration.

The control system for the BASIC airplane had no stability augmentation. This sys-tem is depicted in block diagram form in figure 5. The column, wheel, and pedals were position-command input devices for the horizontal tail, asymmetric spoilers plus ailerons, and rudder, respectively. All servoactuators were assumed to have first-order lags with 0.1-second time constants.

The control system for the BASIC + SAS configuration included a "basic" stability augmentation system (SAS) in addition to the AUTOSPEED. The SAS consisted of longi-tudinal augmentation as indicated in figure 6 and lateral -directional augmentation as indi -cated in figure 7. Generation of the 3 and ay feedback signals is described in refer- ence 4. It is recognized that the /3 signal is not readily available on most real airplanes

11

but with a modest onboard computational capability it can be generated. The important feature of the a feedback is that it consists of both an outer loop ay and 'an inner loop Y6r• By using the gains specified in reference 4, the overall transfer function of the rudder servo became 10 , whereas all other servos were modeled with 10

s+23.2 s+10

The control system for the FULLY AUGMENTED airplane included longitudinal and lateral command augmentation. The longitudinal command augmentation is shown in the top half of figure 8. This system provided well-damped responses to pilot inputs with the column or pitch trim button and inherent pitch attitude stability (that is, "pitch hold") when no pilot inputs were made. The lateral command augmentation is shown in the bottom half of figure 8. This system (or control mode) is described as "roll-rate wheel steering" and provides a slightly stable spiral mode rather than "roll-hold" upon removal of a wheel command. Zero roll rate is commanded whenever there is no wheel input; thus, the system inherently acts to control (reduce) large rolling moments 'due to the failure of an engine. Thus, the task of recovery from an engine failure is made easier for the pilot because he does not have to react instantly to preclude a large lateral transient.

The pitch-attitude command system (fig. '8(a)) in effect replaces the longitudinal augmentation of figure 6 and the roll-rate command system (fig. 8(b)) replaces the lateral augmentation (that is, the part leading into the "aileron servo") depicted in figure 7. The directional SAS (or upper part of fig. 7) is then incorporated to complete the control sys-tem model for the FULLY AUGMENTED airplane.

DESCRIPTION OF SIMULATION EQUIPMENT

A block diagram of the moving-base RDS/STOL simulation system is given in fig-ure 9. Real-time digital simulation techniques were used wherein the digital computer accepted inputs from the pilot's controls and made outputs to the simulation hardware 32 times each second. As indicated in the diagram, the outputs went to (1) the cQckpit instruments, (2) a six-degree-of-freedom servodriven TV camera viewing a scale model of a STOL airport, (3) a remote abstract flight-director display, and (4) an analog com-puter which was used to generate special drive signals for the RDS motion system. A. discussion of this special drive circuitry is presented in appendix A.

The RDS part of the simulation system was a large six-degree-of-freedom test bed located inside an aircraft hangar at Langley Research Center. Figure 10 is a photograph of this hardware. An example of the simulator response is shown in figure 11 which is an amplitude-phase plot of the response in the lateral degree of freedom. The relation-ships shown in this figure were typical of the simulator response in the other axes. A

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general description of the RDS is given in reference 6 and a tabulation of its motion cap-abilities, both normal and as modified for this study, appears in appendix A.

Figure 12 is a photograph of the simulator instrument display and cockpit controls used in this study. The primary instrument 7 in the display group is the attitude indica-tor combined with a flight director. A conventional cross-pointer type of director (see appendix B) was used but the command bars were driven by signals and logic from the main computer program. A block-diagram indication of the flight-director drive signals is given in figure 13. The transfer functions were developed and the gains were opti-mized during the reference 4 study. In the present study only gains K22 and K23 were changed from the reference 4 value. Figure 14 is a photograph of the cockpit dis-plays and includes the virtual image presentation of the airport scene and the heads-up flight-director display. The airport scene is controlled by the same computer as the airplane and thus "grows" realistically as the aircraft approaches the runway. Further discussion of the RDS/STOL cockpit, cockpit instruments, and heads-up flight-director display is given in appendix B.

As indicated in the Introduction, the same equations of motion, control systems, aircraft dynamics, and simulated atmospheric conditions as used for reference 4 were used in the present study. A summary of the equations is presented in appendix C.

RESULTS AND DISCUSSION

Two types of results are presented herein: (1) pilot evaluation of the handling qualities of an externally blown flap STOL aircraft while recovering from an "engine-out" (EO) during landing approach and (2) trajectory errors due to the EO. The approaches began from level flight (trimmed for the high-lift configuration) at an altitude of 243.8 m (800 ft) and about 3048 m (10 000 It) from the runway threshold. A lateral offset of 121.9 m (400 It) from the localizer was also used. The 6 0 glide slope was intercepted about 2320 m (7612 ft) from the threshold. The pilots were instructed to maintain the glide slope by holding a fixed vehicle pitch attitude and varying the flight path by increas-ing or decreasing thrust. Either the left or right outboard engine was failed at some alti-tude between 38.1 to 152.4 m (125 to 500 It) during the approach. (The pilot was usually not warned in advance and a number of "non EO" approaches were randomly mixed in.) The pilots were not given the option to go around.

Time histories of several of the lateral-directional variables are shown in fig-ure 15 for an Instrument Flight Rules(IFR) approach in turbulence with the FULLY AUGMENTED airplane. (The Y-trace is off-scale for the first 17 seconds because of the previously mentioned initial lateral offset.) The outboard engine on the left wing was failed at approximately 60 seconds. The reaction of the SAS was immediate and caused

13

the ailerons to satuate (Oa-trace) and the spoiler on the right wing to activate. The pilot quickly superimposed a large control wheel (o) input which drove the spoilers to a large angle and reversed the rolling moment. This reversal was quick enough to cause the spoiler to deactivate and the ailerons to come off their stops. Three smaller reverse -O inputs were then made to stabilize the roll attitude and limit the transient lateral travel to less than 9 m (30 ft). The ailerons resumed their saturated position and the spoiler operated around a deflection level of about 10 0 to 150 until touchdown. Note that the return overshoot of the localizer (Y-trace) was about as large as the EO transient displacement, but the pilot reacquired the localizer and simultaneously brought the roll angle to zero in sufficient time to make a good center-of-the-runway landing.

Three aircraft-control configurations were flown with and without motion during the EQ studies. A virtual-image TV scene of the airport was used during approaches under simulated visual flight rules (VFR). Automatic speed control (AUTOSPEED) and a flight director were used and were considered to be valuable aids for achieving good con-trol and maintaining the pilot workload at an acceptable level during both VFR and IFR landing approaches. These aids were particularly important (and necessary) when command-control augmentation was not used.

EQ Recovery During IFR Approaches

A summary of the pilot ratings of the handling qualities during IFR recovery attempts under several simulated conditions for each of the three configurations is given in table IV. The Cooper-Harper rating scale was used and is shown in table V. The rat-ings reflect the following:

(1) Only the FULLY AUGMENTED airplane handled satisfactorily under all approach conditions considered. It handled about as well in an EQ situation as during a normal 60 glide-slope approach (in both the motion and nonmotion runs). The reason was that the augmentation precluded development of large upsetting moments which could lead to pilot-induced oscillations (PIQ).

(2) The other two configurations handled poorly anytime the lateral axis was upset. For example, with the BASIC airplane the failure of an engine and/or the presence of turbulence usually led to serious P10 situations, some of which would have been cata-strophic in real life. The P10 situations were sometimes aggravated by the motion of the cockpit particularly when P10 conditions occurred in the pitch and roll axes simul-taneously. Further investigation determined that the motion was valid for initial accel-erations due to a disturbance on the aircraft; but if several sizable oscillations occurred, the phasing between the angular and linear accelerations became unrealistic and the pilots experienced difficulty in determining what their control response to the combined cue should be. The phasing problem was attributed in part to the different response

14

characteristics of the RDS linear and angular drive systems. Consequently, the unaug-mented (that is, BASIC configuration) was removed from further consideration, partly because of its extremely poor lateral-directional handling qualities, but primarily because the RDS was found to be inadequate to accommodate some of its required motions because of a lateral upset.

(3) All three pilots said that it was easier to make the simulated EO recoveries in a nonmotion environment, but they all agreed that the motion environment felt more realis -tic (subject to the preceding PlO discussion) and typical of the conditions they would expect in a real situation. They said that the initial cockpit motion following an engine failure provided a helpful cue and alerted them to initiate a recovery action quickly. One pilot also commented that the simulator motion enhanced the turbulence simulation (which had seemed to be unreal during the fixed-base runs). Nevertheless, in a motion—no-motion comparison (see table IV), they rated the piloting tasks about equal for the FULLY AUGMENTED aircraft and favored the no-motion situation by about one rating point for the BASIC + SAS configuration.

In general, the pilots were not bothered by the failure of an outboard engine during approach with the FULLY AUGMENTED airplane because very little motion was involved. Because of the high-gain attitude stabilization, the pilots had to train themselves to remain out of the control loop until discrete inputs were definitely required. Even then, relatively small wheel and throttle inputs were sufficient and the pilots had to guard con-stantly against overresponding to an unexpected engine failure.

Everything was not back to a normal landing situation after recovery from an engine failure, however, because only a marginal amount of thrust was left for the flare and landing. Thus, there was an increased need to reach the flare point with favorable veloc-ity components and aircraft attitude and to initiate the flare maneuver promptly. In order to accomplish this, the pilots preferred to concentrate strongly on their instruments dur - ing this phase and they tried to ignore all motion cues. The pilots commented that such a technique was not unusual because they also tended to ignore motion cues during the final phase of the landing approach with many airplanes in the real world.

Statistical results of the trajectory dispersions during 40 EO recoveries made with the FULLY AUGMENTED airplane under IFR conditions are shown in table VI. The mea-sures of EO recovery are: (1) maximum altitude loss (1h)EO with respect to the glide-slope beam and (2) maximum lateral deviation ()Eo from the localizer during the transient motion. The point of touchdown (X and Y components with respect to the runway threshold and the center line, respectively) and the rate of sink ii at flare initi-ation and at touchdown are also given in table VI. The ii at flare initiation is included to indicate how well the pilot was able to restabilize the aircraft along the 6 0 glide slope

15

following an EO. (Nominal h on the 60 glide slope is 4 rn/sec (13.23 ft/see) for the 75-knot approach speed.)

The results in smooth air were slightly better for the fixed-base set of runs than for the moving-base runs. The average maximum excursions and the standard deviations for recovery are smaller and the h at flare initiation is closer to nominal with a noticeably smaller standard deviation. The touchdown results are also better for the fixed-base runs. The X-values for the designated landing zone are 76 m to 213 m (250 ft to 700 ft). -

The results of the 10 recoveries and touchdowns with motion and turbulence were the poorest of the three sets, although good control was maintained. (See fig. 15.) The engine failure was detected about as quickly as in smooth air and recovery action was prompt; however, precise reestablishment on the localizer and glide slope was not quickly achieved and the pilots were often still working on small error corrections when it came time for the flare. The pilots initiated the flares at the proper altitude but lift. variations due to appreciable spoiler activity to control roll oscillations during the flare often resulted in relatively hard (mean h = 2.21 rn/sec (7.25 ft/see)) touchdowns. The associated standard deviations were also high because of the turbulence and because there were more roll corrections during the flare in some landings than in others.

An obvious modification to the flare maneuver for three-engine landings in turbu-lence is to initiate the flare at a slightly higher altitude than when all four engines are operational. To illustrate the technique, flares were initiated about 1.52 m (5 ft) higher than the nominal 16.15 m (53 ft) in several additional EO tests. Touchdowns occurred near the end of the landing zone at vertical velocities from 0.95 to 1.75 rn/sec (3.08 to 5.76 ft/see). Thus, it appeared that ideally this STOL could be landed with three engines in moderate turbulence about as well as with four engines in the same turbulence, but the control situation was potentially more critical because of the thrust limitation. Unlike four-engine flares where a thrust program was followed, the procedure for three-engine flares was inherently simplified in that maximum thrust (with three engines) was coinci-dentally the approximate amount required for flaring, provided h was near its hominal value at the flare-initiation altitude. Thus, the pilot usually commanded full throttles to command maximum powered lift (for the three operating engines) at the flare point and then waited for the touchdown. He had little time or capability to alter the flare trajec-tory along the way (unless pitch-command inputs were also made with the column).

EO Recovery During VFR Approaches

Landing approaches were made with the FULLY AUGMENTED configuration under simulated VFR conditions. When an engine was failed, the subsequent recovery was often less hurried than for IFR conditions because the pilot was visually assured that he was

16

maintaining adequate control of the airplane during the EO transient and recovery. Fur-thermore, with the airport in sight, he had an instinctive feel for the magnitude of the recovery action required and the time available to realine the airplane with the runway. Consequently, the maximum excursions away from the localizer and glide slope were somewhat greater than those for IFR approaches, but the recoveries were no less successful.

The pilots said that the difficulty of the piloting task during EO recovery was not significantly changed when the visual scene was added and used. The pilots further stated that the cues from the visual scene were realistic during the early approach, EO transients, and recovery because the gross altitude cues contained in the television pic-ture were adequate for performing the control tasks. However, they were highly critical of the televised scene as the aircraft neared the runway (below 30 m (100 ft)). They said that during the flare maneuver they were not able to obtain the altitude and altitude -rate cues necessary for making good landings. (This deficiency had been noted in previous studies involving TV displays.) A large digital display of altitude was then superimposed on the landing scene, but it was generally ignored; the pilots did not particularly need this type of information above an altitude of 30 m (100 ft) and below 30 m they did not have time to monitor it effectively as they set up for and executed the flare.

When a simple "heads-up" flight-director display (see fig. 14) was superimposed on the landing scene and used by the pilots, they made the recoveries more quickly. The pilots commented that any slight improvement in recovery time could be attributed to their detection of the engine failure a fraction of a second sooner because of the initial visual and motion cues. They further stated that after detection they concentrated almost entirely on the heads -up flight director display until the recovery was nearly complete; then, they started to seek and use other visual references such as the runway. (No determination was made as to the value, if any, of visual cues obtained from the underlying landing scene while concentrating on the heads-up director.)

The VFR part of the study was terminated after it was demonstrated that no unexpected control problems were encountered during EO recoveries with the FULLY AUGMENTED aircraft and that there was no trend toward improved handling qualities because of the presence of the motion and visual cues.

CONCLUDING REMARKS

A moving-base simulation program has been conducted to study the problem of recovery and direct landing of an externally blown flap STOL aircraft after failure of an outboard engine during the final approach. The study results indicate the following:

kvi

1. Rather sophisticated stability and control augmentation systems are required to assist the pilot in maintaining precise control of this type of airplane after failure of an outboard engine. Considerable augmentation will also be required for normal four-engine approaches and landings.

2. The simulator motion and visual cues, in general, contributed realism to the simulation but did not effect the anticipated decrease in difficulty of the piloting task and/or reduction in stability augmentation system (SAS) complexity.

3. As the degree of control augmentation is reduced, serious pilot-induced oscilla-tions (PlO) could result from lateral upsets because of strong turbulence or the failure of an engine. For such aircraft disturbances the initial motions of the simulator enhanced the control situation, but if a PlO tendency arose, then the phasing between the linear and angular cues degraded the motion cues. This condition, in turn, aggravated the PlO ten-dency and caused a degradation in overall control performance. The phasing problem was attributed to the inability to compensate adequately for inherent lag characteristics of the simulator in strong oscillatory situations.

4. The ratings assigned to the piloting task by the pilots during IFR approaches were not changed when a visual landing scene and/or a heads-up flight director were added to the simulation.

5. The piloting techniques and EO recovery results were about the same whether the pilots used the aircraft's regular flight director exclusively or the visual display which included the simple heads-up flight director. However, when the heads-up direc-tor was not included in the out-the-window display, the engine out (EO) recoveries were performed more slowly and maximum trajectory errors were larger, but the recoveries were equally successful.

6. With FULL AUGMENTATION and a reasonable amount of piloting compensation, the subject STOL aircraft could be landed satisfactorily in moderate turbulence with only three engines operating, but the control situation was potentially critical. Because of the thrust limitation, the important requisite for a good three-engine landing in turbulence was being able to restabilize the aircraft precisely on the glide slope before flare initia-tion and then promptly adding full thrust (for three engines) at the selected flare altitude. Inititation of the flare slightly above the selected flare altitude would be desirable if run-way length were not critical.

Langley Research Center National Aeronautics and Space Administration Hampton, Va. 23665 July 2, 1975.

18

APPENDIX A

REAL-TIME DYNAMIC SIMULATOR (RDS)

The Langley real-time dynamic simulator (RDS) is a moving base simulator with independent motion drives in each of its six degrees of freedom. It consists of a simu-lator cockpit mounted in a three-axis gimbal system which is attached to a horseshoe-shaped support frame, which, in turn, is suspended by eight cables from an overhead carriage-dolly system. (See fig. 10.) The cables are arranged in a manner such that they provide nonpendulous suspension of the cockpit for horizontal accelerations less than 0.25g. The gimbals provide the three rotational degrees of freedom: the inner gimbal axis providing roll; the middle gimbal axis, yaw; and the outer gimbal axis, pitch. The three translational degrees of freedom are obtained through the carriage-dolly sys-tem. Vertical motion is achieved by rotating a large drum which winds up the eight cables attached to the gimbal support frame. The cable drum is mounted on the dolly which rides on tracks mounted across the main carriage. The main carriage moves along another pair of rails (see fig. 10) attached under the roof of an aircraft hangar at Langley Research Center. The simulator cockpit for the present study was oriented as shown in figure 10. The landing approach was assumed to be in a direction approxi-mately normal to the direction of the main carriage motion. This arrangement permitted use of the large operating envelope of the main carriage for the potentially large lateral motions associated with failure of an outboard engine.

The basic RDS drive systems are capable of the following maximum displacements, velocities, and accelerations:

RDS motion drive performance

Cockpit axis Maximum displacement Maximum velocity Maximum acceleration

Lateral ±22.86 m (±75 ft) ±2.44 rn/sec (±8 ft/sec) ±2.44 rn/sec 2 (±8 ft/sec2)

Longitudinal ±2.44 m (±8 ft) ±1.22 rn/sec (±4 ft/see) ±1.22 rn/sec 2 (±4 ft/see2)

Vertical ±6.1 m (±20 ft) ±1.83 rn/sec (±6 ft/see) ±1.83 rn/sec 2 (±6 ft/see2)

Pitch ± Continuous rotation ±1 rad/sec ±1 rad/sec2

Yaw ± Continuous rotation ±1 rad/sec 11 rad/sec2

Roll ± Continuous rotation ±1 rad/sec ±1 rad/sec2

19

APPENDIX A

However, the motion was restricted to the following values for the present study:

Restricted motion drive performance

Cockpit axis Maximum displacement Maximum velocity Maximum acceleration

Lateral ±22.86 m (±75 ft) ±2.13 m/sec (±7 ft/sec) ±2.44 m/sec 2 (±8 ft/sec2) Longitudinal 0.0 (0.0) 0.0 (0.0) 0.0 (0.0)

Vertical ±6.1 m (±20 ft) ±1.83 m/sec (±6 ft/sec) ±1.83 rn/sec2 (±6 ft/sec2)

Pitch ±0.7 rad ±1 rad/sec ±1 rad/sec2

Yaw ±0.35 rad ±1 rad/sec ±1 rad/sec2

Roll ±0.35 rad ±1 rad/sec ±1 rad/sec2

The vertical and lateral motion drives were not restricted, but the longitudinal axis motion drive was deactivated completely because significant longitudinal accelerations were not expected. The restricted angular motion was dictated by the size and shape of the STOL cockpit. The use of these restricted motion envelopes, however, did not inter-fere with the motion requirements for the present study.

The RDS drive components in each axis are, by necessity, large and fairly complex. Hence, system response is dramatically affected by the interaction of the various mechan -ical subsystems comprising the total drive system. Therefore, during the initial design and construction of the RDS facility, internal electronic circuitry was used to minimize these interactions and to optimize the system response over a fairly large frequency range. Hence, when operating in a restricted motion envelope, the drive signals can be tuned for best response in the restricted region by preconditioning those signals. To accomplish this for the present study, an analog computer was inserted between the digital-computer model and the RDS. (See fig. 9.) The analog computer was used to manipulate the drive signals to obtain the best dynamic response of the RDS, with respect to the model, for the frequency range required for the present study. A combination of several motion drive techniques including "motion scaling," "washout," "feed forward," and "motion onset" were used. The resultant computer-generated motion-drive signals reproduced motion cues, within the capability of the simulator, that were most nearly representative of those produced by the simulated STOL aircraft. The following compu-ter schematic diagrams show the circuitry used to manipulate the drive signals for each axis. Below each diagram is its associated transfer function in Laplace transform representation.

Yoror Z'

or Dr '

APPENDIX A

Lateral and vertical:

• A(Cs2+s)

Y (s+B)(s2+Ds+E)

Pitch and roll:

A[E(s-i-D)-C] 9(1+ EB)(s+D) -CB

A[Es2 + (ED + C)s]

s2+(D+BE)s+B(C+ED)

Yaw:

21

APPENDIX A

In addition to manipulating the drive signals based on prior theoretical analysis, once the circuitry was developed, experienced NASA research pilots tuned the final sys-tem by performing various maneuvers with the simulated aircraft and had the perceived acceleration levels (and other angular cues) adjusted to match the dynamics of the com-puted aircraft as indicated by the cockpit instruments.

As is explained in the body of the report, the motion drive was Only partly success - ful. Good correspondence was obtained for most motions, but, on occasion, when an exaggerated PlO tendency occurred, the induced motion would degrade rather than enhance the simulation.

22

APPENDIX B

THE SIMULATOR COCKPIT AND VISUAL DISPLAY

The cockpit which was installed on the Langley real-time dynamic simulator (RDS) for the STOL landing approach simulation was essentially the left half of a side-by-side four-engine transport airplane cockpit. The interior of the cockpit (attitude controls, instrument package, and various cockpit items) was arranged similar to the fixed-base simulator cockpit used in the reference 4 study.

A pilot's view of the instrument panel and the flight control components is illus-trated in figure 12. Only the components identified by number were used in the present simulation. The primary instrument monitored was the "flight director" (designated 7).

A close-up view of the flight director is given in figure 16. The components of the flight director pertinent to the STOL simulation fall into three basic groups: attitude, situation, and command. The attitude group consisted of the horizon line, pitch- and roll-attitude sphere, miniature airplane symbol, and roll-attitude pointer. The situation group consisted of a glide-slope pointer, localizer pointer, and the radio altitude bar. The command group consisted of a pair of flight-director bars driven by a signal com-posed of the elements and shaping indicated in figure 13. Acquisition of and staying on the glide slope and localizer were accomplished by manipulating the appropriate vehicle controls so that the flight director bars were kept centered on the dot in the middle of the miniature airplane symbol. The preferred procedure for acquiring and maintaining the glide slope during a landing approach was to hold a fixed-vehicle pitch attitude and vary the flight path by increasing or decreasing thrust. The horizontal situation indicator (HSI) located directly below the flight director provided heading information and duplicated the localizer information presented on the flight director.

The various lights and switches are identified in figure 12. The three lights located over the altimeter and numbered 9, 10, and 11 were used to indicate warning and initiation cues to the pilot during the simulated landing flare and touchdown. The vertically arranged pair of lights (designated 16) were used to indicate to the pilot the proper trim position for the throttles (designated 18) prior to initiating a test approach. Individual instrument illumination was provided by installing post lights at the airspeed, altimeter, rate-of-climb, angle -of -attack, and direct lift control (DLC) instruments.

The switches located on the control wheel were as follows: On the left horn of the wheel a button (designated 1) was used to engage the autospeed control. Usually the pilot activated the autospeed circuit with this switch immediately after beginning a test approach. The thumb switch (designated 2) on the left control horn was a four-way air-craft trim switch for pilot-initiated changes in vehicle pitch and roll trim. The DLC

23

APPENDIX B

thumb wheel (designated 14) on the right horn was an optional method for making small changes in thrust. This option was generally ignored in lieu of the throttle.

The control column and wheel dimensions are shown in the following sketch. The control motion dampers were set for a control damping ratio of 0.7. They were discon-nected during the system force-deflection calibrations.

Adjustable roll control damper —a

(13 in.) H33.02 cmH

- ..—Hand grip center line

Roll-force Adjustable springs

pitch control damper -.

Pitch-force spring

A force-deflection diagram for pitch control is shown in figure 17. A similar dia-gram for roll control is shown in figure 18. The control forces chosen were a composite of forces from military specifications (ref. 7), pilot opinion (NASA test pilots), and vari-ous aircraft of similar configuration and performance. (For example, see ref. 8.) A roll-control force-gradient of 0.4 lb/deg was chosen (that is, 0.2 lb/deg/hand). The pitch-control force gradient for these tests was 14.23 N/deg (3.2 lb/deg).

Rudder pedals were installed in the cockpit. For the FULLY AUGMENTED air-plane rudder pedal inputs were not required. The pilots did use the rudder pedals when the degree of augmentation was reduced.

For VFR approaches, a virtual image scene of a typical STOL airport was pre-sented to the pilot. The airport scene was obtained by using the computed aircraft motions to drive a six-degree-of-freedom color TV camera system over a 1/300-scale model of a simulated STOL airport. This scene was transmitted to the simulator cock-pit by closed circuit TV and displayed on a 63.50-cm (25-in.) color TV monitor which was mounted on a bracket forward of the windshield. The pilot viewed the scene on the monitor through a virtual image lens which represented the windshield. (See fig. 14.)

Thrust-command information for glide-slope tracking and roll-command informa-tion for localizer tracking were combined into a heads-up flight director display located

APPENDIX B

appropriately in the windshield field of view. This display consisted of two fixed parallel bars and a moving dot which appeared to be focused at infinity rather than on the wind-shield itself. A typical location used during the study is shown in figure 14. The dot was usually zeroed hallway between the right-hand ends of the parallel bars (as shown in fig. 14). One pilot preferred a single-line display with the dot zeroed at the midpoint of the line.

The bar-dot display was generated on a cathode-ray tube (CRT) and viewed by a second TV camera. The video signal from this camera was then mixed with the video signal from the camera viewing the airport scene. The combined signal was transmitted to the 63.50-cm (25-in.) monitor at the simulator cockpit.

25

APPENDIX C

SIMULATION EQUATIONS

The aircraft equations of motion and related equations are presented in this appen-dix. The forces and moments acting on the aircraft are summed in a conventional body-

axis system XB, yB' z with origin at the center of gravity of the aircraft. The gravity force and wind conditions are defined in an Earth-fixed or inertial axis system X1, Y11 Z 1 and transformed into the body system.

The linear acceleration components in the XB, YB, and Z B directions, respec-tively, are given by

PVA ru1 [

-g sin 9 + yr - wq + 2m Cç 1 I

= PVA2S PVASb g cos 9 sin + wp - ur + 2m Cy

' + 4m C2 (1)

UM g cos 9 cos +uq-+ C I PVA2S 2m

CX = CX,o + Cxo ot + Cx6* 63 + Cx0 sp0sp + CX0 IOsI + Cx, ge + CX,El (2) f3 f

C = Cy + Cyô r 0r + C a sYJ 6a + C Os + Cy El (3) •

= Cyp + Cyrr (4)

C Z = C z ,0 + C 0 O + C z63 O '3 + C 0 + C 6 IOsI + C z, ge + C Z , El (5)

where

C = C ,1 + C,2

The coefficients on the right are, in turn, defined analytically in the section "Symbols" and numerically by the values in table H. The spoiler deflections (see fig. 2) are always upward but the asymmetric-spoiler deflection angle ös is defined in terms of

26

APPENDIX C

the rolling moment it induces; thus, 0s appears as an absolute value in equations (2) and (5) but not in equation (3) because the sense of the Y force is dependent on the sign of Os. Coefficients with the subscript "El" are used with appropriate logic to introduce the effects of failing either outboard engine.

The angular acceleration components are given by

_i ply - Iz — qr+ 'X

--Z (j.

-i;+ pq) + PVA2Sb C1,1 + PVASb2 C1,2

I Ix 21 41X

q Iz - Ix pr+ (r2 - p2) +PVA2 C +

m,1PVA2 Cm,

I ly ly My My

rJ IX 'XZ( pq+— - qr) +PVA2Sb

C + ,1

PVASb2Cn, 2

- Iz I 41z

(6)

where 1x' ly, and IZ are moments of inertia about the body axes of the aircraft and IXZ is the associated product of inertia. The nondimensional moment coefficients on the right of equation (6) are defined by

C1,1 = C1130 + ClörOr + C10 a + C10 o s + LCl,El (7)

C1 ,2 = Ci + Clrr (8)

Cm, 1 = Cm , o + Cm0 Ot + Cm03 O 3 + Cmo sp O sp + Cm 0 1 0 s1 + CM, (9)

Cm , 2 = Cm q + CmJ1 (10)

C11 1 = Cn 13 + Cn0 r 0 r + Cn0 a 0a + Cn0

s O s + ACn,E1 (11)

C ,2 = Cp + Cnrr (12)

27

Ug U

Vg = ajj v

Wg W

(i=1,2,3; j=1,2,3) (is)

APPENDIX C

where

C1 = C1,1 +

Cm = Cm , 1 + Cm,2

Cn = Cn,i + Cn,2

Specific values of the coefficients on the right-hand sides of equations (2) to (5) and (7) to (12) were determined from table II by means of a "table-lookup" interpolation scheme in the computer program.

The wind conditions used in the simulation were included as follows. The total aerodynamic velocity of the aircraft is given by

VA(UA 2 +vA

2 +WA2) 1/2(13)

where

tiA U+Ug

VA = v+vg(14)

WA w+wg

The components Ug, Vg, and w are transformed wind velocities given by

where u, v, and w are defined in the Earth-fixed axis system as constants or as functions of altitude. The direction-cosine elements of the transformation matrix [au] are

28

APPENDIX C

all = cos 0 cos q.' (16)

a12 = COS 0 sin V.' (17)

a 13 = -Sfl 0 (18)

a21 = sin 4) cos V.' sin - cos 4) sin p (19)

a22 = sin 4) Sifl V.' sin 6 + cos (P cos V.' (20)

a23 = sin (P COS 6 (21)

a31 = cos (P Cos V.' sin 6 + sin (P sin V.' (22)

a32 = Cos 4) sin V' sin 6 - sin(P cos V.' (23)

33 = COS 0 COS 6 (24)

K'4 The Euler angles V.', 0, and (P are derived from

r cos 4) + q sin (P

cos 6

= q cos 4) - r sin 4) (25)

4)

=p + V.' sin 0

In addition to the wind conditions just discussed, random turbulence could be super-imposed on the velocity components in each axis. The turbulence model is based on the Dryden spectral form and involves feeding Gaussian white noise from three independent random number generators through shaping filters. The filter transfer functions are presented and discussed in reference 4. Rotational gusts were not simulated.

The origin of the inertial axes XI, Y1, and Z1 was located at the runway thres-hold, with X1 pointing down the runway center line and Y1 to the right. Altitude h is calculated as -Z1. Equations for the inertial displacements are determined from

29

APPENDIX C

• -1 X

Y=aij V

-li w

(i = 1, 2, 3; j = 1, 2, 3) (26)

where [au] -1 is the inverse of the [aij] used in equation (15). The glide slope was assumed to intersect the runway at X = 76.2 m (250 ft) and the localizer source was assumed to be at X = 1219.2 m (4000 ft). The landing zone was designated as 76 X 213 m (250 X < 700 ft). The angles E and e are aircraft position errors with respect to the localizer and glide slope, respectively. They are determined by

tan =

(27) 4000 - X

Ez = tan-1 h - 12 - Glide-slope angle (28) 250 - X

The instantaneous altitude h in the ez equation was biased 3.7 m (12 it) because the landing gear of the simulated aircraft extends this distance below its center- of gravity. (The variables X, Y, and h must be expressed in feet when used in equations (27) and (28).)

REFERENCES

1. Crane, Harold L.; Sommer, Robert W.; and Healy, Frederick M.: Effects of Reduced Airspeed for Landing Approach on Flying Qualities of a Large Jet Transport Equipped With Powered Lift. NASA TN D-4804, 1968.

2. Parlett, Lysle P.; Greer, H. Douglas; Henderson, Robert L.; and Carter, C. Robert: Wind-Tunnel Investigation of an External-Flow Jet-Flap Transport Configuration Having Full-Span Triple-Slotted Flaps. NASA TN D-6391, 1971.

3. Grafton, Sue B.; Parlett, Lysle P.; and Smith, Charles C., Jr.: Dynamic Stability Derivatives of a Jet Transport Configuration With High Thrust-Weight Ratio and an Externally Blown Jet Flap. NASA TN D-6440, 1971.

4. Grantham, William D.; Nguyen, Luat T.; Patton, James M., Jr.; Deal, Perry L.; Champine, Robert A.; and Carter, C. Robert: Fixed-Base Simulator Study of an Externally Blown Flap STOL Transport Airplane During Approach and Landing. NASA TN D-6898, 1972.

5. Vogler, Raymond D.: Wind-Tunnel Investigation of a Four-Engine Externally Blowing Jet-Flap STOL Airplane Model. NASA TN D-7034, 1970.

6. Connor, Andrew B.; Bergeron, Hugh P.; and Schoonover, W. Elliott, Jr.: Ride-Quality Research Activities at NASA Langley Research Center. Symposium on Vehicle Ride Quality, NASA TM X-2620, 1972, pp. 229-246.

7. Anon.: Flying Qualities of Piloted Airplanes. Mil . Specif. MIL-F-8785B(AG), Aug. 7, 1969.

8. Innis, Robert C.; Holzhauser, Curt A.; and Quigley, Hervey C.: Airworthiness Consi-derations for STOL Aircraft. NASA TN D-5594, 1970.

31

TABLE I. - RANGE OF CONTROL-SURFACE DEFLECTIONS AND

MAXIMUM DEFLECTION RATES

Control surface Range of deflection angle, deg

Maximum deflection rate, deg/sec

Horizontal tail ±10 50

Rear wing flap 0 to 90 5

Spoilers 0 to 60 50

Aileron ±20 50

Rudder ±40 50

32

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0 0 0 0 0 0 0 0 0 0'''''l''U,

0 0 0 0 0 0 0 0 0 ''C'

0 0 0 0 0 0 0 0 0

.

CO CU 0 0) U, U, CU U, 0 •••CUU, COO,U,CU00I

CUCOOCOCUCUCUCU 0 C- U,COCOU,NCOU,U, CU 0 0 0 0 0 0 0 0 0 0 II COU CUCUCUCOOCU 0 0 C- 000000000 0 0 0 0 0 0 0 0 0 a 0000000000 0 0 0 0 0 0 0 0 0 CUU,U,U,U,U,CflU,CU 0 0 0 0 0 0 0 0 0

Es d's.-sod ,,,,,, •4-d•,•..•

00 O0)U, N OOCO CU 0 0 0 CU U, 0 N COCOCUCUCUCUCUCO CU CU 0 CU 0 0 00 CU CU 0)OCU 0 CUCO 00 0000 eCU U,O U, 0 U U

Es C, 0OO' 000000 0000000 0000000 U q,•''.•

- — U, o .0U,0U,0U,0U, 0 ' ' CU CU CU 00)00)00)005 0

5 ' SICOCUCUCU 0U,0U,0U,0U, 0 C .5 - CU CU CU — 0U,0U,0U,0U, 0 ' ' CO CU CU CU

33

N EU

- weou,00CCc,e

EUNNOU,U,COCC

NONOCCOCCOCO 0 o - — NCC0O

U

d EU

UNNOCCOON

IC ONO ONNU,US EU 0

N

EU01CC

0 -. UCCU,O0CN00

EU

N CC CO N U, CC CO CC U, 0 - >< ONOOCC-- EUU 0CC CC

0

EUOCCCC

0,1CC

- COCC00000CCOO N

N — 0 U, 0 CC U, 0 N CO N CC CC - N CO CC N CC CC CC CC CC CC CO 0 CC 0 0 0 0 0 0 - N —

Ii U$NNC0CC 000000000 EU 0IICCCCCC 0

C -

0 U,CCOCCNOCO N

CC N CC 0 CCC CC N CC N - N 0 O CO N 00 N CC NNNNNNN 0 0 0 0 e 0 — - - 000000000

0 — 0 CC N N 0 - N 0 N

EU o CCCCCCCC 0 C

NCC000NNOO 0 II

0 — N N N N — 0 0 0 0 0 0 0 CC 0 - — — N 0 - N. 0 0 0 0 0 0 N 0 0

EU 000 0000 000000000 o 01 001 CC 0 — ON OCCCC ON 00

CC U, — CC CC N U, CC 0 o o 0 0 0 0 e e

CO 0 0U,U,U,N 0 0 0 0 0 0 0 0 0 0

EU 0 CC CCCC 0

- NNNO ONNO o —

NCCNCCU,ONNU,N N 0 0 0 0 0 CC 0 N

U, 00 0 00 — 0 N 0 0 0 0 0 0 0 0 0 C 00 IR IR

EU o

0 CCCCCCII CC Cd

0 •CICCCCI

0 -0 0 0 0 0 0 0 0 0

0 N 0 00 0 N 0 CC

0 0 0 0 0 0 0 0 0 Cd

EU 0C CICCCC0ICCCC C

000NON 00NN N U

d ii N 0 0 0 0 0 0 0 0 0 0

0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 EU OC CICCIC 0

0

-

N U, N — 0 - — - 00 N N 88888

o 0 0 0 0 0 C I 0 C C I

CC CC'CCCC 0N U,NN • U,U,,NOU,CC 0 0 0 0 0 0 0 — 0

888 CL 8 8888 001 0

U, U, U, U, N 0 U, CO CC 0 0 0 0 0 — 0 •

888888888 8888 88 8 0 0 ..C0

- •0 .UCCU,NCOCCONO O888888888 ct 88 8888-8

EU U e CC 000 .

.

N CCNU,NOCCNNCC 00000N0 Ii 888888888 888888888

EU 0 0

0 — — 0 N 00 N CC 00 CC N 0 N — — 0 0 0 0 N N 0

88888 888888888 U 0000 C

-CCIII III

0.U,0U,0U,0U,0 0U,0U,0U,0U,0 Cd I — — N N N I — N N N

z 0 EU

0U

Cd, C Cd

0

z

34

TABLE II.- BASIC AERODYNAMIC INPUTS USED IN SIMULATION - Continued

(b) Spread engines. (Inputs same as for clustered engines except for coefficients herein.)

C,,=O C T=1.40C T 2.81 C T=O CT= 1.40 CT=1.40CT=2.81 deg _C T =O

Cy , E1 C1-1,E1

-10 -0.010 0.065 -0.017 -0.002 -0.040 -0.060 -0.042 -0.135 -0.190

-5 0 .045 -.029 -.002 -.040 -.069 -.020 -.130 -.192

O -.005 .032 -.035 -.003 -.038 -.073 -.003 -.129 -.200

5 0 .048 .003 -.001 -.041 -.083 .002 -.146 -.223

10 0 .055 .025 -.001 -.038 -.088 0 -.161 -.251 15 0 .050 .037 -.001 -.044 -.095 .002 -.185 -.278

20 .015 .070 .125 -.002 -.049 -.102 .002 -.178 -.318

25 .015 .185 .225 0 -.061 -.114 .002 -.253 -.392

30 .020 .075 .175 -.004 -.036 -.077 0 -.133 -.270

cy , per (leg C0 per (leg Cl , per deg ba y 6a

-10 0.0005 -0.0058 -0.0042 0.0008 0.0002 -0.0010 0.0024 0.0038 0.0041

-5 -.0025 -.0022 -.0050 .0008 .0002 .0002 .0013 .0035 .0046

0 -.0032 -.0051 -.0068 .0009 .0002 .0008 .0003 .0034 .0049

5 -.0055 -.0074 -.0082 .0008 .0006 .0010 -.0001 .0038 .0056

10 -.0050 -.0078 -.0088 .0008 .0004 .0014 0 .0041 .0058

15 -.0050 -.0075 -.0094 .0010 .0006 .0018 .0020 .0052 .0068

20 -.0062 -.0085 -.0134 .0018 .0006 .0020 .0044 .0046 .0084

25 -.0062 -.0112 -.0182 .0018 .0019 .0021 .0045 .0116 .0120

30 -.0075 -.0068 -.0128 .0016 -.0011 .0004 .0026 .0033 .0108

35

TABLE II. - BASIC AERODYNAMIC INPUTS USED IN SIMULATION - Concluded

(c) Ground effects (same for clustered and spread engines)

h CL = 3.0CL = 3.5CL = 4.0 CL = 3.0 C L = 3.5CL 4.0 CL = 3.0 CL = 3.5 C L = 4.0

M ft LCLge LCD,ge 1Cm,ge

0 0 -0.090 -0.154 -0.239 -0.19 -0.24 -0.29 -0.15 -0.30 -0.48

3 10 -.020 -.050 -.095 -.08 -.11 -.14 -.08 -.15 -.23

6.1 20 -.008 -.026 -.050 0 0 0 0 0 0

9.1 30 -.002 -.005 -.009 0 0 0 0 0 0

Cx, ge = LI CL ,ge 5fl a - LC D, ge cos a

C z,ge = _LCL , ge cos a - iCD, ge Sifl a

a)

±5

Ct) N CO Ca .-. CO CO CO CO Ca 0 CO N CO Ca - Ct) N C) N 0 CL) CC CO CO Ca - CO CO CO Ca

CO 0 Ca N CO Ca N CO CO CD C) CO CO Ca N C) Ca Ca Ca N CO CO Ca CO CO CD CD co N 0) ,-C co CO wC CO V 0 v N 0 Cav U) N CO C) 0 0 - -4 Ca Ca Ca

CO — Ca CO CO CD CD CO N N N N N N N CO CO CO CO CO CO CO

- CO CO CO CO CD N N C) Ca CO C) N CD 0 CO CO CO CD 0 0 N 0 CO 'C CO — CO N

0 Ca CO Ct) 0 Ca N 0 N N CO N Ca CO 0 N eq 0 C) N N CO CO N N C) CO -C Ca Ca ' CO CO CO N 0 Ca CO N CO CO Ca CD C) Ca CO CO CO N

CO Ca Ca Ca Ca CO CO N 0 U) CO C-C CO CO 0 .4 Ca CO CO C CO CO CO CO CD CO CO co C) - - — Ca Ca Ca Ca CO CO CO CO CO CO CO CO CO CO CO CO CO CO CO

N N CO Ca -C CO CO CO CO Ca CO CO CO CO U) Ca CO N Ca Ca 0 N CO 0 -4 Ca CO U) CO N CD CO 0 Ca N CO Ca N CO W CO Ca N Ca N N CO V CO CO Ca CO N Ca V CD CO Ca CO CO CO CO CO CD N Ca C CO CO CO CO N 0 . Ca CO CO CO CO CO N N N N CO CO CO

- ,4 C1 Ca CO CO ' '

Ca ,-C CO CO CO CO CO N N Ca .- CO CO N — N '-4 CO CO N CO Ca CD Ca Ca CO CO CO Ca qC -C 0 Ca CO CO 0 Ca N ' Ca 'C Ca -C N CO 0 CO CO r -C -C 0 Ca CO CO N CO CO C

CO CO N N C) CO 'C Ca Ca ' CO -. N Ca U) 0 ' N Ca C•C C N CO Ca 0 v CO CO

Ca Cl Ca C4 CO v CO N 0 Ca CO CO N CO Ca Ca Ca Ca 0 0 0 0 0

CII . - ,C -C — C ,C .4 ,C .0 C•L Ca Ca Ca Ca Ca Ca Ca Ca Ca Ca

N N CO Ca — CO CO CO CO Ca CO 'C Ca -, CO Ca Ct) N Ca 0 CO CO CO CO CO N CO Ca

CO 0 Ca N CL) Ca N CO w CO CO CO N -. 0 CO 0 w N C) CO CO N Ca ,C CO U) CO CO CO N 0) - CO CO CO Ca a' CO CO Ca Ca 0 0 0 0 — - -. Ca Ca

- .0 Ca Ca CO CO CO CO CO CO ' a' ' 'U'

N .4 Cl) CO CO CO CO N N 0) - Ca Ca Ca CO N N 'UC a' CO 'UC CO C-C CO 'U' C-'U -C 0 Ca CO CO 0 Ca N 'U' (M N CO Ca CO C!) CO Ca C 0 Ca CO N N CO CO 'U CO CONNCaCOCaCaCOCOCOCaCOCOCOO C')COCONCO CO Ca Ca Ca Ca CO 'U' CO N 0 Ca V. CO CO CO C- N N CO CO CO CO CO CO CO CO CO

- - - - - - - - - - - - - - ,-C - - -

'U' N CO Ca -C CO CO CO 'U' 0) N CO ,-C CO CO -C CO 'U' CO N CO Ca 0 -4 Ca Ca CO 'U -C CO 0 Ca N CO Ca N Ca 'U 'C CO -. Ca CO Ca CO CO Ca 'U' CO CO -C C') CO N Ca '-C

CO CO

CO CO CO CD C- Ca -C CO CO CO CO CO CO Ca 0 0 0 ,C .4 ,- ,C Ca Ca Ca CU Ca CO - - Ca Ca CU Ca Ca Ca CO CO CO CO CO CO CO CO CO CO CO CO CO

CO CO CO CO N 0 0) CO CO 'U CO CO CO CO CO CL) CO Ca CL) Ca Ca CO CO CO - 'U -C 0 Ca CO CL) 0 Ca Ca 'U' Ca 'UC 0 N CO CO Ca -C 0 Ca Ca CO N N CD CO 'U N CO N N Ca CO 'C Ca 0 a' -C Ca CO CO .-C 'U' CO N Ca Ca 0 -C Ca CO 'U' U) CO N 'U' ca Ca Ca Ca CO 'U' CO N 0 -C - Ca Ca CO CO CO CO CO CO 'U' 'U' 'U' 'U' 'U 'U' 'U' 'U'

- - - - - - - - - - — - - - - — — - -

"-C Ca N CO C) ,- CO CO CO Ca Ca CO CO N C) 0 . Ca 0 Ct) CD 0 Ca N CO Ca N CO Ca CO Ca 'U CO CO CO

CO CO CD CO CO N C) ,-C Ca CO CO 'U' 'U' 'U' CO CO CO

CO CL) CO CO N 0 0 " CO N CO N 0 'U'

Z0 Ca

- 0 C) CO CO 0 Ca 0 CO CO 'U' CO" '-' C 0 CO N N Ca CO 'C Ca CO CO -C CO 'U' CO N CD Ca

CO Ca Ca Ca Ca CO 'U' CO CO CL) CO CO CO CO CO CO CO

C.)

CO C- C- C) Ca CO CO Ca CD N C) CO CL) 'U' CO CO CO CO 0 Ca 0 0 CO N 0 Ca -C Ca CO Ca 'U' CO Ca CO CO CO 'U' 'U' 'U' CO CO CD N C) Ca CO CO N N CO Cl CO 0 CO CO CO CO Ca 0 0 'C Ca Ca Cl CO -C - Ca CO 'U' CO CO CO N N N N N N CO CO CO CO CO CO

0'U' C -C CO CO CL) 'U' -C 'C CO CL) CO 0 - CO 0 CO 'U' Ca N .44 Ca Ca CO Cr 0 CO N N CO • 'U'

a) CD CO CO N CO CO 'U' CO '-C CO 'U' CO 0 - CO ,4 N CO -C CD CO N Ca CO ,C N 'U' -C Ca N N CO Ca N CO CO N CO 0 -C CO CO Ca CO CO CO Ca CO CO Ca CO CO 'CI' 'C 'U' 'U' -C N Ca CO 0 Ca CO CO N

o CO .-4 'C -C -C Ca Ca Ca Ca Ca CO 'U' CO CO -. CO - 'U' N Ca -C Ca CO 'U' 'U' CO CO CO CO CO CO CO

0. CO -C - Ca Ca Ca Ca CO CO CO CO CO CO CO CO CO CO CO CO

Ca a) - CO CD Ca 0 CU Ca 'U' N Ni:l CO Ca CO Ca -C CD 0 CO Ca CO N 0 CO 'U CO N Ca 0 -C CO CO CD N 0

U) C- N Ca CU CO CO Ca CO N Ca CO CO 'U' CO CO -C N N N CO 'U' ,-C N Ca C') U) Ca Ca 'U' CD Ca 'U' CO CO CO

Ca CO CO CO 'U' 'U' 'U' CL) CO CO N Ca Ca CO,CO 'U' COO -C Ca CO 'U' CO.CO CO CO CD CO N N N CO CD CO Ca C)

0 C, CO CC -C CL) CO CO 'U' -. -C CO CL) CO 0 -C CO CO CO 'U 'U' N ,"C -C Ca CO Ca CO Ca CO Ca CO CO CO Ca 0 CO

Ca COCD N CO CO 'U' CO -C CO 'U' co 0 -C CO Ca CO 0 N C') 0 CO CO 'U' CO Ca " 0 C) CD N CO CO 'U' CO Ca

E- 0 co. to N COO CO CO Ca CO CO CO Ca CO CO Ca ' COO 'U' NO CO 'U' CO N CaCaO Ca 'U' CO CO Ca 0

Ca ,C -C .0 Ca Ca Ca Ca Cl CO 'U' CO CO ,-C CO to co CO C) C) Ca 0000 000 'C -C -C Ca CU ,.0 ,C -C ,- ,-4 ,4 Ca Ca Ca Ca Ca Ca Ca Ca Ca Ca Ca Ca Ca Ca

CO CD CO C) 0 Ca Ca 'U' N ,-C N CO Ca CD Ca Ca CO 'U' 0 'U' N Ca 0 Ca CO 'U' CO CO N C-

N N C) Ca CO CO Ca CO C- Ca CO CO 'U' CO CO 0 -C 'U' CU CO Ca CO Ca C CO CO N CO CO CO 'U' 'U' 'U' CO CO CO N C) Ca CO CO CD CO 'U' CO CO CD CO CD N N N N

CO Ca Ca CO CO CO CO CO CO CO CO CO CO CO CO

CO , ,4 CO CO CO 'U —C - CO CO CO 0 ,C CO CO Ca -C Ca CO CO CO CO CO CO 0 CO CO Ca CO CO NCO CO 'U' CO CO 'U' COO ,C CO CO 'U' ON CO 'U' CO Ca -C ,C0 Ca N CO CO N CO 0 ,C CO CO C) CO CO CO Ca CO CO CO N CO CO Ca ,C Ca 'U' CO CO C- N

CO ell CaCaCaCaCOCOCOCaCOCOCOCOCOCDCOCDCOCOCO

CO CO

CD CO C) 0 Ca Ca 'U' N 'C N CO Ca CO 0 Ca N U) C) CO CO Ca -C C- Ca CO CO CU CO N Ca CO CO 'U' 'U' Ca -C CO CO CO Ca N Ca N Ca

0 CO CO CO 'U' 'U' 'U' CO CO CO N Ca Ca CO , CO CO CO N CO CO C) 0 0 CaCaCa CaCaCa CaCOCO

CO ,C -C CO CO CO 'U' ,C C CO CO CD 0 -C Ca 0 CO CO N C CD - CO CO

CU CO CO N CO CO 'U' CO —C CO 'U' CO 0 ,C CO Ca 'U' ,.'C C) N CO CO Ca CO CO CO N CO 0 CO CO C) CO CO CO Ca CO CO 'C Ca CO CO CO Ca 'U' CO

Ca Ca Ca Ca Ca CO 'U' CO CD Ca' 0 -- Ca Ca Ca CO CO CO

CD CO CO Ca 0 Ca Ca 'U' N ,C C- Ca Ca CO CaCO CO N CO CO N N Ca Ca CO CD Ca co N Ca 0 'U' CO CO 'U' 0

O , CO CO CO 'U' C , CO CO Ca CO 0 0 'U' 0 N 0 CO CO

to N CO CO 'U CO , CO 'U' - CO 'U' 0 N CO CO CO CO CO N CO 0 .0 CO CO Ca CO CD CO 0 CO Ca Ca 'U' CO

CO Ca Ca Ca Ca Ca CO 'U' 'U' CO CO CO CD CO CO

N CO CO Ca 0 Ca Ca 'U' CO CO N CO CO

N C- Ca Ca CO CO Ca 'U' CO CO CO CO CO 'U' 'U' 'U' CO CL) CL) U)

CO CO CO U' 'U' CO ,C

COco CON CO CO 'U' CO C11 ,C CO CO N CO 0 , CO 'U' CL) CO

Ca Ca Ca Ca Ca Ca Ca —

C.) Ca 'U' CD CO 0 Ca CO CO 0 Ca CD CD 0 Ca 'U' COOR0 Ca CO CO 0 Ca 'U' CD 0 Ca CO

CO U' U'U)COCOCOCOCOCDCOCDCO 0

ri ra

CO 0

CC

Ca 0

0 C

a)

a)

a)

Ca

37

0 0 C Q)

a) - a

Q) 0. <

U) 0 U)

0

0

a) U

0

0 0

1.

Lj

U) U) en a eq en U) en U) eq

U) U) N - U) en en U) c' - en N 0 en eq

U) en U) N N N U) U) U)

N U) eq N U) en .

N a - 0 eqw 0 U) U) oa en co eq N

co N - en U) to U) eq eq en en en en en en

In N eq en 0 ' eq U) en N - en U) eq

en en ta a U) co U) - U) U) U) U) N en In N U) 0 a en eq

U) In U) N N N N U) U) U) U)

q. en U) eq 0 N N en N en U) N

en N N U) eq U) . U) U) N a a q. N In 0 U) a en N U) a a C13 en ' en n U) U) U) en eq eq en en en en en en en en en

a U)

N U) eq . U) U) U) a en a N en en U) en U)

- en

U) U) U) a a a en en ' . . en en en en en

en en U) U) en 0 N a en en U)

en en - 0 U) N N en N U) en en q U) N U)

eq a en eq C4 C4eq eq eq eq eq eq eq eq eq eq eq

In U) - eq en en a eq U) 'q. U) en U) en en U) eq

N eq en U) U) 0 In U) 0 eq en en eq U) N en eq co en en U) U) U) a en eq cq U) en n en U) U) N N N N N U) U) U) U)

co -. en co eq en N a a U) n

N U) eq N U) 0 N U) - N en 0 N U) N eq a U) U) 1 en eq N eq U) eq en N

U) U) U) q. N a eq en en en U) U) CO en en - C11 C13 en en en en en en en en en en

a U) en en - U) en U) 0 U) en

N eq N a en U) U) N en U) N U) a a en en en . en en en

en U) en eq U) U) N N en

Ien U) eq co en U) N In

c- eq en a U) en U) U) - eq U) U) 0 en en eq eq eq

eq - en eq eq eq eq eq cq en U) eq U) a en en

.0 a N - en a U) 0 I N - eq en en q.

zU)

U) a N N U) U) q eq 0 U)

en N eq oa en q. en U) U) U) U) U) U) U) -

III

ee

U)NU)enenU)U)U)eqU)enen0enen ) U) eq

U) N en Nw U) en U) U) N en U) eq U) N en N a U) In U) eq co N U) U) a en eq eq

U) eqenenU)enU)U)NNNNNNU)U)U) 0 0. ' N U) to N U) en U) en U) U) N N N 0 N U)

Q)to N U) U) U) eq en U) U) en U) cq U) en en U) N CO

en U) N en en en U) eq eq U) en a q U) eq en N U) eq en U) eq co U) a eq en en q en en en U) U) to en en en en eq eq eq en en en en en en en en en en en

en en N - U) N U) eq en U) en 0 en N en N - en en to U) U) U) 0

' q. en eq en en

N CD

N U) (0 en N eq eq en a a N ( 7) U) U) 0 a U) en en U)

en en U) N C, q en U) U) q en q eq M- U) U) a a en en eq eq U)

en _en COQ eq eq eq eq eq -

U) U) en U) en eq en q. en U) eq

ento U) N eq U) U) a eq N U) U) a a e a .-I en

en eq eq eq en en en en en en

In N U) '-I U) en en U) eq en

•a U)

en N en en en eq en en a en co eq . en U) N U) U)

en eq eq en en enen en en en

In N U) U) U) en N a a U) n1 a q U) en U) eq en (0 eq

en U) eq en en U) U) U) a U) U) eq a eq en en 0 q' en .n N In U) en eq N 0 en en N U) U) a en eq eq eq

U) eq eq en en SS N N N N N N N U) U) U) U) U)

U) N en U) v en co e U) a a . a a I I en ' U) N U) en N C,3 en U) en en N en U) en N en a U) N CO .1 en U) In U) U) eq en C', N en N eq co eq v CO N U) en eq In a r N 0 en C4 en en . en en U) U) (0 U) en en en en C4 C11 C11 en en en en en en en en en en en en en

en N U) eq 0 U) a eq en U) U) a en U) Q

en (0 N q a en U) a U) eq . N U) en N eq I to N U) U) U) 0 0 a a eq eq en en en en en en en

U) U) N 10 eq U) a a N N N en I U) U) en

U) en en-0 ..1 N en en en en 0U) en en a en a U) en en CO U) a en en

U) en eq q. .n U) U) N N N N U) U) U)

en N en U) q. o en U) en a U) U) U) U) 1 en U) q U) en CO U) U) en eq en en U) eq

0 oa In U) U) eq en U) en en U) en co en N 0 U) N en N en en U) a eq U) a a eq en coen a a N U) 0 eq en en U) U) U) U) a a en en en eq eq eq eq U) enenenen C,3eq . (1 en U) U) N N N N N N N N U) U) U) U) U) SSS U)

U) en a a U) co N N en en en en CO U) en N N N n a . v U) en q. CO N N 10 0 CCen N en U) en en eq U) U) n I U) en en eq en a U) U) N N U)

. U) a en U) 01 en N en q eq co q. en N a q N U) en eq env en U) N U) I (JI en N U) eq N eq en U) a en eq eq en 0 en n en to U) U) U) U) U) U) en en en eq Cq eq en en en en en en en en en en en en en en en en en en -

C) eq CO U) a eq q. U) U) 0 eq U) U) 0 eq Ir U) U) a eq U) U) 0 eq '

In a

38

5 ri

Cl)

C

Cl C

0 o Cl

U) z oC

U) c

C Cl

zCl

Cl Cl

------------

39

C) C.) 4-Cl,

0

0

ca

C)

0 Cl)

'O C 0 0 0.

0

a) C a)

a) C.. 0.

Ew r LJ

C) N - C) U) eq C) N U) cq N U) CC eq 0 -4 C) C) 0 N .4 U) () C.3 CD 0 CC C) U) C) U) C) 0 C) eq C) N CC eq '-4 0 0 C) C) C) N N CO CO U) U) U) V p ' C.) CC

- - - -

C.) CC C.) CC CC eq eq U) C) C) , 'N q' eq C) U) '-4 C) .4 C.) U) N C) '4 CCv U) N C) eq CC v CO CO N C) CO ' ,-1 co U) eq 0 N U) CC 0 co CO v C.) .4 C) C) N CO

U) U) '.)' .)' CC CC CC CC C.) C') C.) C.) C.) -4--4

C) C) U) . N C' 0 CO CC 0 C) U) CC eq 0 C) C.) N - CC ')' CO C) C) CC N C.) CO ' C.) 0 C) CO C.) C) CO CO U) .)' C.) C') '4 'l 0 0 C) C) C) C) C) C) C)

eq ' 0 N N v 0 0 0 CC CC C) CO eq C) U) eq CC CO C) '4 C.) CO C) eq ' CO N C) C) 0 '4 eq CC

0 CO CC 0 N U) C') .4 0 C) C) N - C.) N N CO CO CO U) U) U) CC C.) C.)

C) 4 ,-4 '-4 '-4 C') C') ' CO 0 U) .4 N C.) C) U) eq C) CO CC 4 C) N U) CC C') 0 0) C) N C.) .4 0 C) co N 0 CC '4 co 0-'-4 C.) v CD 0 CC N '-4 N C.) C) v 0 CO C.) C) L-m 0 C) CO CC .-4 C) C) CO U) C.) CC C') '-4 0 C) C) C) N N CO CO U) U) U) .)' C.) C.)

CC C') e.) eq C') .4 .4 .4 4 .4 4 -4 '

, C) 0 CO CO eq U) N N -4 N CC 0 CO CC eq c Cq C.) U) C) 4 ' N U) C) U) '4 C) CC 0 N C.) 0 N C') N CO C) '4 ' CO C) 1 C.) Lo N C) C) 'i

"w U) CO N C)

CO 'W U) U) CO N N 0 C.) N .4 N V 0 (0 eq 0 N V -4 M N CO V eq 0 C) N CO

'C C.) eq .-4 0 C) C) C) N CO CO CC) U) U) v v v C.) C.) C.) eq C') C.) C') C') C.)

C.) N CC 0 C) CO CC U) ' '4 CC U) v CO C. U) -4 0 C) N CO U) C) -4 0 0 C) C) C) C) C) C) co eq CC CC CC C.) eq C.) C.) eq eq C')

C) U) U) C) ..) C.) C) CO '4 C) C) eq .)' U) N 'C N co C.) C) CO C) CO CC '4 C) C) co N N CO eq C') C.) C.) CC C') C.) eq eq C') C.)

C) C) '-4 .4 N CC eq ' CO C) C') N CC C) N CC 0 N U) CC -4 C)

.0CC co

C) CC

CC eq eq eq 0 N CC v C) ' .4 0 co

N CO

U) .)' v U) .4 CC N 0 (0 eq C) CC U) v CC eq C.) -4 0 0 C) C) co co

CC CC CC eq eq eq eq -4 '4 '4 .4 .4 '4 'l 4 '

Z eq U) C.) co 0 C.) CO C) 'l v v N v 0 CC 0 0 C) C.) CO C) C.) C') .-4 N CC C.) eq 0 U) .4 CO 0 CC N 0 .4 vCO N C) 0 '4 CC N .-1 CC v.-4 CO N C) eq C) v C) CO ')' 0 N ')' C') . C) N CC U) v co C') 0 C) co C) N CO CO U) U) U) U) CC CC

.4C N 0 0 '4 0 C) CO U) C.) V CO C) C) U) C) C.) C) C') CO C.) C) N (0 0 CO C.) ' C) CO U) v CC Cq 0 0 C) C) C) C) N N CO (0

- - C 0 ')' C) ,-4 eq C) CC CC eq C) 0 COC.) U) U) U) C) co C) 0 N v0. CO C.) ' ' '-4 N ' C) eq ' co 0CC-')' CO C) C) - CC U) U) CO Ci) 'I' C) 0 N co C) C) CC C) ') 0 N co '-4 co U) CC '-4 C) CO U) ')'

NNCOCCeq.-4,-4OCOC)C)C)C)C)C)C)NNNN

- ., N 0 C) N it) ')' ')' C) C) CO C) CO CO LI) N C) 0 CO C) CC C) CC N C.) C) CO eq 0 CO eq C) t-CC 0 CO eq C) CO CC 0 C) N CC) ')' CC eq . 0 0 C) C) C) N N CO CO U) U) LI) U) ')' v C.) CC

')' CC U) C) N C) eq C) C) 0 0 N 0 N C.) C) C.) 0) "U.) U) U) C) -4 C) '4 N n U) .4 C) C') eq U) U) CC) v C) CC C) CC CO 0 eq Ct) N C) .4 CC ')' CO C) C) .4 -4 CC CC U) CO N C) CO C) ')' U) -4 N ')' CC C.) C) 0 U) Co CO eqw ')' C') C) N U) C') C) N CO U) C.) C') 0 C) N CO

'l N CO ')' C.) '-4 0 C) C) N N CO CO CC) U) ')' V CC CC C.) e C') C') C.) eq C') C.) - 'l -4

C) CO C') CO C') U) 0 CO CC 'f 0 C) N 'i eq CO N 0 0 - U) - C) N CO C') C) U) C') -4 0 C)C) co N CO eq ')' CC CC CC CC CC eq eq eq C') eq

N ') 0 eq U) eq U) '-4 '-4 ') 0 N 'I U) U) CO (C) .4 .)' N C) 0 '4 ' C) N ')' C) it) C) C.) C) CO U) ')' C)

C) N U) ')' C.) CC eq C') eq eq 4

C) 'l C) CO ')' C') Q C) C.) C') N C) U) ,4 N C') o U) eq -4 -4 4 0 0

N CC) ')' U) CD .4 ')' N '4 U) CC CO N C) 0 '4 C) 'I 0 (0 ')' eq '-4 C) N 0 C) C) C) C) C) N 'l eq . .4 '4 4

C) -4 N 0 0 '-4 Ct) CC ')' CO C) eq CO eq C) U) -4 N ')' -4 C) CO C.) '4 C) CC C') CC CO .01 C) CC eq U) 0 CO CC U) CO C) co ' U) C) ') C) eq C) CC C) U) eq C) ')' 0 CO '1. eq 0 C) N CO U) ')' CC CC eq .4 0 0 C) C) C) C)

')' ')' CC C') C') eq eq C') eq ' '4 .4 'l ' 4 -4 -4-4- -4

eq Ct) N 0 C'.) C) CC C.) CC U) C') C) N ')' 0 '-4 N CC CC CC CO CO CO C) eq C.) U) ')' N 0 CC ')' eq C) ')' C) CC N Cl CC CO C) -4 CC U) CO C) 0 .4 C.) N C) '-4 CC ')' C) C) C) ' ' C) eq C) U) -4 N ')' '-4 C) CO CC -4 C) N CC 0 C) N U) C') ,-4 0 C) C) C) N N CO CO CO U) U) U) ')' ')' ')' ')' CC CC

C.) '-4 -4 4 'l .4 .4

C) eq CO C) 0 eq ')' CO C) 0 eq CO C) 0 eq ')4 CO 0 eq ')' CO C) 0 eq '1' CO C) 0

0 ••• )C.)eqeqCCCCCCCCCC C)LC)UCC

40

TABLE IV. - PILOT RATINGS OF AIRCRAFT HANDLING QUALITIES

ON LANDING APPROACH (PRIOR TO FLARE)

Control configuration Motion Turbulence EO Average* pilot rating

FULLY AUGMENTED 2 (Spread engines) X 2

X 3 X X 3

X 2

X. X 3

X X 21

X X X 3

BASIC + SAS 4 (Clustered engines) X 4

X 4

X X 6.

X 4.

X X 5 X X 6 X X X 7

BASIC 7 (Clustered engines) 10

x 10 X 8 x x •10 x x 10.

*Average for the three pilots.

41

cli CO 0) 0

U) a)

... 0) '

0

U)

C Ca Ca 0 U) C

.2 IC

Ca CU_

a)

CI) .

0

0) 0. 0. 0)

0. . 0.) 0.) E 1U)

.2.0 C

0

a)Q .

C .0 0 0.10 o

9.20.1-.U)

1.. •'U)

.2 Ca 0. .a) 0 dC 0), 0)0 0)

.

C.. 0)0

0 Ca

•C U) 0) a) Ca C.0 0 ,o

). Ca -a a '0 a)'0 U)01 0 C-. 00 ..Ca rl + C ..2

.0 Ca

Ca a) .9 U) C

UUCUE 0 -.

a)'0CC

..-. U) 0. 0

Ca8

a) C0. COU)0. )Ca0 .00)co U)

U) - 0 00) '. I.

C 0

U)U) Ca) . I.. 0 C.. 0

'0U) '0 C U)

.

'0 U) C C's

Ca 0,9 C -0 ' 00 0)0 C

Ca... C 00) 0.00 C C . C U) 'b U) 100C.) 0 Ca0 0.0

EQ 0)Ca U) 0 U) 0

U) 'O0.• O -

C 0) Ca 0)0 0) U) ,9C

I. a oa)

.Cd

0.0 0U)

a) C.)

'0 0 C-. 0

0)

9 Ca -'.2.2 Ca0 C

U) C

I

90

a)

00)

d.,C00

C•

0

U0a)

E. rn Q Ca C.

- a)... E40)0

0

E . 0)

Ca0a)

C

Z0.

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U)

C Ca.00 .00.0)

r1a'0 00

CaCaE C..

0

U) b00) Ca

0 .0C 0

Z

oCa I.. C 9. '0

0)

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.0 0)0.C 0)0

0

CC (D (L'

1.0)0)QQ

o

C

. . 0 0) U)

.Cd. C Ca 0,

. z

E-.C a)

...o.-. .2 U) 0

0 'o

- sCaC) U) U) Ca..-. '0.U)

Z U)

0 ' 0) 0.0)0 .0 .0

cl .0 1.. Ca

C..

'0..-.0 0)00)

1.a)Ca

00

.-0) z o 0.0

CaU)0

Cd

0

42

r.4

0 0 pq

0 0 CI) F'

CI)

F'

C.) o 00 a) a) a) (I) U) (a U) U)

-- — - Cl) - . - -

.IH- I cq 2 .- N c cc co N cc .-iv cc CV)

a) E, L() L) N C) C) (M 0 to cq cq Co L()

C' CO '-4 N tO) CV) C4 r4 CV)

0 m co

bD00 C) C) a) a) (D a)

o U) U) U) CO) — —..-----

Cl) -4-i 41 -11 4-4

- 4.J 4 -.

-6-1 ..

4-i 4-

4-i 4 4-• $-

4J 4

U) Co . C) co Co Co 0 tO) 0 0 C

tO) N C C) CO l CO to) Ci r4 C '-I 0 N C '-4

00 00 Cd a) a) a) a)

U) U) U) U) --.---

-EE

CV) 0 tO) C) v CO C• 1 N Co CV) LoCV) C) C —I Co tO) 0 N Co

•

CO to)

{ CO L') CV) N LO ,-4 CV) LO tO) '-I

C.) 0 00 CU a) a) a) U) U) U) U)

Z14 Z+4 U) I Q 0 CO C) 0 CV) C) .4

cc

ca Co -4 eq Co —I - Cd '-1 C4 - C' CV) cq C 0 tO) .-1

00 00 i

cda) a) a) a) CO Cl) U) CO —- -.--------

0 qw 0 LO r N tO) Co CV) -I tO) —1 CV) v C) CV) cc CV) tO) cq tO) co

OQNCo CV)

'-I

a) 0

a) 0 •- 0 .- 0 .- 0 0

0 Cda) + •• a)

.- '4-4

Cd -Ik'

Cd •-

-' .' -4 -' 4J k Cd rd 0 cd cd

Cd 0 cd 0 cd z cd 0a) Cd a)cd a)cd a) Cd a)Cd

4a)cd

o ,C/) CJ) Cn4. Cl)

4. CI) cd CI)

4

a) 0 —

0 0 N

VII

VII

"-4

0 LO cli

cq

CV) a) .l U)

VII

Ci

.4 CV) VN

VII CV) —4

CIQ C.) • a) Co

N

0 N

a) 0

p. 0

o0

0

Cd

Cd

a) .4 -4

--4 co a) 0

Cd .0

43

'4

Pz Ili

at

ci)

Cd

U)

I ci)

'-4

Cd

w

'-4

co

4. 0

cv

ci)

ci) ci)

'-I

a)

0 0 H

IJ\

'.0

C

Lr\ -(\J

H

4-) .0

ci)

o C'i 0\ NJ N' c aD - - H H CO C\J - N- N- NJ co - - aD

• H .-D '.0 H C

co NJ N--

CO 10 LrN C'.) N- N- (D

+' 0 (N N- 0\ (N 0 H '.0 '.0 '.0

Q) 4.) 0 H - tf\ N-

() (N NJ NJ a) (N (N '.0

0

0 0 0

r4 4.)

C cci

1- 0

0 0 — — — C') .0 H NJ NJ NJ +

C'J 0 4.) 4.) 4) Ct-c

4.) 4- CH I

+ 0 -P I 1)0

-1 .r4 H10

Cci) 'o 1) f)

0 0 0 H

NJ cci cii H H H U)

0 0 0 U) c/) I') 1)0

I NJ cci 0 0 C44 NJ C') NJ 0 ci) cci 0 0 0 0 I

P4 a) I I I Ccci Cd UI ccl Cc)) 1)0 1)0 .)

C]) .0 Ccl) 1)0 0 0 0 ccl 0 r1 r4 C]) Ci) '< >-c N-I <

C) }-4 -4- I-i HI

44

C')

4)

0 -' 0

Ha) U) a)

ri to

a)

1-. Ca

r

Co

H

OlCa I-I

HICa H 1,1 CaIi Ca I

10 IIH

-. .

H r L . — ri

00 I Cd

0 I () I I a) a) I

Cia)

0 0)1

01 I 0

I a) ci

bD0, a)

H

a) Cd

Ia)

—

•0

0 - ' -C a)

Ci) olr

01 cd

b1Cia)

0

CD

—4

a) H ri

0

a)

04

Ca

H

0 C Co

0 C

0 in N

Co

cq

to

-.4

to

45

Touchdown

6,

20 L 7Engine lost

-20

32

deg 0

-32

0 10 20 30 4Q 50

Time, sec

Figure 3.- Typical response of ailerons and spoilers to failure of outboard engine on right wing.

V H Filter

10

_______ S S+10

VAH Filter

Figure 4.- Automatic speed control (AUTOSPEED).

66

46

6

Li

ow

6

op

Engine response

0th character Thrust

(TABLE 111)

Figure 5.- Basic flight controls.

47

6

AT

of 3

asp

at

Figure 6.- Basic longitudinal augmentation.

48

a

Figure 7.- Basic lateral-directional augmentation.

49

oc

ÔLT

6

-e

(a) Pitch-attitude command system (o denotes pitch angle error and AOc denotes commanded pitch angle change).

am n +1 Aileron Gain integrat6w --^ k_^ servo H 0, 6a

Lhnterconnects

(b)Roll-rate command system (p c denotes commanded roll rate).

Figure 8. - Command augmentation systems.

50

>-

-J —u-

LLJ

—_J

Q

< LULU

LL Q

C-) LU >< -J C_) 0 w 0

zE

-- -J w

- c..Z I -J 0

V•i - 01 < I

zL/ ()

LU0

(0

Ln Z Oc

= r)H

<I

[LJ

UJ

I-

LUUJ

I-0 -J

(r) LU >

0

0

in

Q

uj

-J

I-

LU = C-)

0 I-0

E a) U)

U)

0

Cl)

Cl)

a) b.O

L ---J

51

L-75-181 Figure 10. - Real-time dynamic simulator (RDS) with STOL cockpit mounted

in the gimbals.

52

co H

N--1

cq

'-I L

0)

H H OOw0)

t.o a) o o .,-i ,O 0 0 H

.,-1 + + H C, i-4 r-1 U) C.) •r-1 -rl 0)

H H + a) 4i I -1 O+++ a)+

C) ow 0+) U) (i) a) r1 0 +' -c 05 r1 H ,O 0

(0 --7T IJ '0 N- CO H H H H -1 H

005 H 0

H rd

•rI + 0

HO tto

+' 0 +' H +' H O H 054) bO

05 H 0 '0 H 0+' 0+' HC OS H 1c0 +' b

c) 0 oS H • •,-IHH

H 0+) H 05 H

050o U) 0 o H 0 +'

+'O 'da)

1)OH -i -c OH H cci cci 0 H OHH OH

N- co O 0 H (\J HHH

00 H 0Pc

O+'4-' H O 0 H +'r-la) 00) +'HH OS a) 0Oc +''d ,Oa) +'H

'0 050) rd rd II) C rd CH a) O505 0)00 Pc Ha) a) c" Pc a) (L) 00 Wr - IH +) 4^ cfl$. 0 H -H 'H (\J (fl UN \D

0

a)

C.) 0 C.)

4.

0

cii)

U)

4-. 0 —4

'-4 U)

54

0 0) 0 ñ a)

CD '00 C'i c .Jd , .—•1

II II II

C\J C"J çf• C)

0)

CVJ

C.) 0

F

±

csJv

.N ca E

00 —JO

+

- a)_ 0•) cc E

1 a) n — — —

00

— — ± ± n

I-., (-•

c'J cJ

VIA > > Ui U)

'L) —& a)

1 — II II

00

E0 Z a) 0

- M a) 0 CA

• (f\. LA cl_s

CD CQ

IILt\

II II II

c'.i Lr\ 0 N-

I- I-

a) 0

a) r-en

0 CD U

U)

Cd

(I)

a)

Cd ii

C.)

a) r 0 0

a) — C

bD

cJ '—I

a)

0

-NJ

U)

> Ui

55

'-4

bD Cd

U)

U)

ci) '-4 ci)

I 0 LI, —

L) co

4-. 0

bD

U)

ci)

56

50

Control wheel deflection,

O, deg 0

-50 25

Aileron deflection,

•0a' deg0

-25 40

Spoiler deflection,

°s deg 0

-40 10

Angle of roll, ,deg 0

-10 40

Angle of heading,

qi,deg 0

-40 125

Lateral displacement,

Y,m 0

-125I I I I I I I I I I I I I 0 20 40 60 80 100 120

Time, seconds

Figure 15.- A typical time history of the lateral-directional variables during an approach involving an engine failure.

57

58

=) 0—

CD

0 Co L)

—o 00 —

-I IC)

Th E E 0 L)

E E 0 L)

0 rvl

L a

.4-

C

C

C

.4-

4-4.

(1

C

0

Cd

C)

--4

I

-16 -12 -8 -4 0 4

Column forward Degrees

Figure 17.- Column force-deflection curve.

8 IZ 16 Column back

40

160

30120

2080

10 40

z

0 C..) L.) 0 0 0

-10 -40

-80 -20

-120 -30

-160

-40

-200

59

4-

Ol Of

00

CD

CD

a)

C.) CM 0) 0

U,G)

Q)

a) c\J C.)

0

-4 a)

CD

'-4 CD

bb

[4

CD

0

CD0)

CD CD CD CD CD CD CD CJ 00 00 C\i — I I —

N 'DJO

q i 'auoj

60 NASA-iang1ey,'1975 L-9743

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