A flight investigation of the stability, control, and handling qualities of an augmented jet flap STOL airplane

NASA Technical Paper 1254

of

A Flight Investigation

the Stability, Control, .

and Handling Qualities of

an Augmented Jet Flap ,STOL Airplane

Richard E . Vomaske, Robert C. Innis, Brian E. Swan,andSeth W. Grossmith JUNE 1978 \ I TECH LIBRARY KAFB, NM 0334354

NASA Technical Paper 1254

A Flight Investigation of

the Stability, Control,

and HandlingQualities of

anAugmented Jet Flap STOL Airplane

Richard F. Vomaske and Robert C. Innis Ames Research Center, M o f e t t Field, California

Brian E . Swan

C a n a d i a nA r m e d Forces, Ottawa,Ontario Seth W. Grossmith CanadianDepartment o f Transport,Ottawa,Ontario National Aeronautics and Space Administration Scientific and Technical Information Office SYMBOLS AX acceleration along longitudinal axis, m/sec2 A Z acceleration along vertical axis, m/sec' b wing span, m

-

C reference aerodynamic chord, m drag drag coefficient, - CD drag coefficient excluding hot thrust contribution cDA CoG dragcoefficientingroundeffect(excluding hotthrustcontribution) cold thrust .___ CJ jet momentum coefficient, isentropic, QS lift CL lift coefficient, qs lift coefficient excluding hot thrust contribution Q A C L ~ liftcoefficientingroundeffect(excluding hotthrustcontribution) Wnz total lift coefficient, cLT qs rolling moment rolling-moment coefficient, . - -~ ~- C I ijSb

acl

- C l i ai pitching moment cm pitching-moment coefficient, QSZ pitching-moment coefficient in ground effect cmG ...

yawing moment yawing-moment coefficient, TSb

a c n

t h s t thrust (hot) coefficient, -

P

drag force drag-force coefficient, body longitudinal axis,

qs

% X side force side-force coefficient,

qs

lift force C Z lift-force coefficient, body vertical axis,

@

CG center of gravity, percent F FcoL controlcolumn(stick)force, N (Ib) F W control wheel force, N (Ib) g acceleration due to gravity, 9.81 m/sec2 (32.1 7 ft/sec2) h radio altitude, m (ft) i aerodynamic coefficient independent variable supersubscripts CJ, c w p 0, or 6 , IX moment of inertia about the longitudinal axis, kg-m2 iv moment of inertia about the lateral axis, kg-m2 moment of inertia about thevertical axis, kg-m2 aerodynamic coefficient independent variable supersubscripts p , q , or r vertical load factor load factor per unitangle of attack, g/rad nZCy high pressure engine rotor rpm, percent (1 00 percent equals 12,135 rpm) NH roll rate, rad/sec P roll acceleration, rad/sec2 P P period of oscillation, sec pitch rate, rad/sec pitch acceleration, rad/sec2 - free-stream dynamic pressure, - pv , N/mZ (lb/ft2) 4 .

r yaw rate, rad/sec i yaw acceleration, rad/sec2 revolutions per minute S Laplace transform variable S wing area, mz (ft2) t time, sec time to 1/2 amplitude, sec 7-1 12 time to double amplitude, sec T2 V velocity, m/sec (ft/sec) equivalent airspeed (EAS), knots VE maximum operating airspeed, knots 'MO VFR visual flight rules V weight, N (lb) W angle of attack of fuselage corrected for upwash, positive nose up, deg angle of sideslip, positive nose left, deg rate of change ofsideslip, deg/sec flight-path angle, positive up, deg aileron deflection, positive T.E. down, deg choke deflection, positive T.E. up, percent (1 00% equals 38") control column deflection in pitch, positive aft, deg elevator deflection, positive T.E. down, deg flap deflection, positive T.E. down relative to the wing chord plane, deg rudder pedal deflection, positive left forward, cm (in.)

rudder deflection, positive T.E. left, deg spoiler deflection, positive T.E. up, deg control wheel deflection, positive clockwise, deg damping ratio pitch angle, positive nose up, deg V nozzle deflection, positive down from full aft, relative to fuselage datum line, deg ambient air density, kg/m3 (slugs/ft3 ) P apparent roll mode time constant, sec rA roll angle, positive right wing down, deg roll angle after 1 sec, deg yaw angle, positive nose right, deg

*

undamped natural frequency, rad/sec vi A FLIGHT ’INVESTIGATION OF THE STABILITY, CONTROL, AND HANDLING QUALITIES OF AN AUGMENTED JET FLAP STOL AIRPLANE Richard F. Vomaske,* Robert C. Innis,* Brian E. Swan,? and Seth W. Grossmith$ SUMMARY Thestability,control,andhandlingqualitiesofanaugmentedjetflapSTOLairplaneare described. The airplane is an extensively modified de Havilland C-8A “Buffalo” military transport.

The modified airplane has two fan-jet engines which provide vectorable thrust and compressed air fortheaugmentorflapandforBoundary-LayerControl (BLC). The airplane is equippedwith powered flight controls, and stability augmentation is provided on the lateral and directional axes.

The test conditions of airplane gross weight varied between 165,000 and 209,000 N (37,000 and 47,000 lb).

Stability, control, and handling qualities are presented for the airspeed range of 40 to 180 knots. With theStabilityAugmentationSystem(SAS)operating,thehandlingqualities are consideredsatisfactoryfor the normaloperationalenvelope(65 to 160 knots airspeed). Without stabilityaugmentation,thelateral-directionalhandlingqualitiesareconsidered marginal in the landing approach. Poor turn coordination and spiral instability are the primarydeficiencieswhen operating with the SAS inoperative.

Thehandlingqualitiesinvestigationincludedoperationwith one engineinoperative(simu- lated). The cross ducting of augmentorandboundary-layercontrol (BLC) air, togetherwith adequate control power, provide acceptable handling qualities for the one-engine-inoperative condi- tion. The powered elevator control system, a modification incorporated in the airplane, enhanced controllability in pitch, particularly in the landing flare, and allowed investigation of airplane stall characteristics.

INTRODUCTION The augmentor jet flap or augmentor wing powered lift concept has been recognized by both government(ref. 1)andindustry(refs. 2-4) as one of the promisingconceptsforresearchand development.Afterextensivewind-tunnelstudies(ref. 5)and simulationstudies(refs. 5-8), a researchaircraftprogram was initiated to provide an aircraftforSTOLflightexperiments.A de Havilland C-8A Buffaloairplane was modified into an augmentor jet flap research aircraft. A description of the modified airplane and a summarization of the airplane contractors development program is presented in reference 6. Important aerodynamic features of the airplane are presented in reference 9.

*Ames Research Center, Moffett Field, Calif. 94035.

?Canadian Armed Forces, Ottawa, Ontario, K1A OK2.

$Canadian Department of Transportation, Ottawa, Ontario, KIA ON8.

I .

This report presents the results of stability and control, and handling qualities flight testing of the airplane. Test results peculiar to the airplane with the spring-tab elevator system are presented in references 10 and 1 1 while thepresentreportincludesdataandinformationpertinent to the airplane with the powered elevator. Because of the limited elevator control authority usable with the spring-tab elevator system, flight tests with the spring-tab elevator controls did not include flight conditions atangles of attackexceeding 2 I".

The stability and control characteristics of the airplane are in some respects unique because of powered lift effects and the slow speed flight capability. The powered lift effects have a primary bearing on the longitudinal stability and control while the slow speed flight capability significantly affects both the longitudinal and lateral-directional stability and control. Piloted simulators (refs.5-8)were used during the design phase of the airplane development to study the unique stability and control characteristics and the aircraft handling qualities.

The initial airworthiness flight-test program was conducted by the contractor and reported in reference 6. During theseteststheairplane was flown withinaflightenvelopeoffrom 50 to 180 knots and at load factors sufficient to demonstrate that the aircraft flight loads were within design and that the airplane was flutter free. Some preliminary documentation of the stability and control characteristics were obtained and reported in reference 10.

Results of the initial research flight tests to document andevaluate the performance, stability, control,andoperationalcharacteristics of the airplane are presented in reference 11. Results presented in reference 1 1 arefortestsperformedbeforetheinstallationofthepoweredelevator controlsystem.Theairworthinessflight-testresultsof thepoweredelevatorcontrolsystem are presented in reference 12. Aerodynamic and operational performance characteristics are presented in references 10 and 1 1.

THE RESEARCH AIRCRAFT The research aircraft is a highly modifiedde Havilland C-8A Buffalomilitaryturboprop transport.Its high "T" tailmadeit especiallysuitable forapplication of apoweredliftsystem.

Photographs of the aircraft in the landing and takeoff configurations are shown in figure 1 . Table 1 lists the geometric and mass characteristics of the aircraft. Figure 2 is a three-view drawing of the aircraft. Special features of the aircraft are described briefly below; a more complete description is 6. Reference 12 presents details of the powered-elevator installation which is presented in reference not included in the description of the airplane given in reference 6.

Engine Two Rolls Royce Spey M K 801-SF split flow engines provide thrust for the airplane as well as air fortheaugmentorflap,body BLC, and aileron BLC. Eachengine is fittedwithvectorable conicalnozzles (fig. 3). The nozzlescanvector thethrustbetween 6" and 104" relative tothe to provide lift and flight-path control. The nozzle vector angle control handles, fuselage centerline, one for each engine, are located in the cockpit overhead console adjacentto the engine throttles and within easy reach of the pilot (fig. 4).

Air Distribution System The air distribution system directs the engine bypass air to the upper and lower augmentor nozzles, to the fuselage boundary-layer blowing nozzles, and to the aileron blowing nozzles (fig. 5 ) .

A crossover ducting system is used so that approximately 64% of the bypass massflowof each engine is ducted to the augmentor and aileron nozzles on the opposite wingand to halfof the fuselage boundary-layer blowing nozzles; the remaining 36% of the bypass mass flow is ducted aft to the augmentor nozzles on the same side of the aircraft as the engine. The air distribution systems foreach engine are completelyseparate,butidentical.Thisuniquearrangement provides for engine-out operation without large rolling o r yawing moments.

Augmented Jet Flap Theflapsare deflected by hydrauliclinearactuatorsmountedexternal to the wing. The minimum (flaps up) angle is 5.6" and the maximum flapangle is 72'.

The ducts that supply the air to the augmentor flap nozzles are mounted just aft of the rear spars and provide air independently from each engine. The inner ductair is supplied from the engine on the same side of the aircraft as the engine, while the outer duct is supplied from the engine on theopposite side (fig. 5). Thearrowsshowninthefigureindicateairflowfrom the right-hand engine only.

Fixed leading-edge slats were installed t o help maintain airflow over the wing at the high values of circulation obtained during powered-lift operation.

Flight-Control System Several modifications were made to the basic C-8A Buffalo flight-control system. In the cockpit (fig. 4), the control wheel was replaced with a wheel instrumented to read out lateral and longitudinal control forces. The only other change in the cockpit flight controls was the addition of an electrical lateral and pitch trim switch, which the pilot could control with his left thumb (fig. 4) and the copilot with his right thumb; a manual elevator trim controller is provided for emergency operation. The engine nozzle controls were installed adjacent tothe engine throttlecontrols (fig. 4(a)).

The longitudinal control for the first 78 hr of flight testing consisted of a spring-tab elevator system.Adescription of thespring-tabsystemandtestresultsarepresentedinreferences 10 and 11. Subsequent flight testing of the airplane was with a poweredelevator system installed. The poweredelevatorwasinstalled t o improvelongitudinalpitchcontrol.Adetaileddescriptionand airworthiness test results of the poweredelevator system are presented in reference 12. Longitudinal pitchcontrol is providedbyasingle elevatorsurface whichprovides both maneuverandtrim control functions as the horizontal stabilizer incidence isfixed.Thefrictionandbreakoutforce characteristics of theexistingsystemarepresentedinfigure6.Thefeelsystemincorporatesa computer that varies feel gradient as a function of dynamic pressure, as shown in figure 7. The feel gradient is constant for speeds less than 60 knots. Control column to elevator gearing is shown in 8 and control authority and trim limits are shown in figure 9. The nose-up elevator authority figure decreases with increasing airspeed because of the available aft column travel and a reduction of the power control unit steadystate gain. This reduction in authority and gain is caused by the flexibility of the power control system components.

Thedirectionalcontrol consistsofatwo-panel rudder,theaft panelbeinghinged tothe trailing edge of the forwardpanelandgeared to it in a 2: 1 deflection ratio. The rudder isfully poweredthroughanirreversibledualhydraulic actuatorcontrolledby cables fromtherudder pedals. The basic Buffalo directional control system was modified t o include a stability augmenta- tion system (SAS) actuator which is mechanically summed in series with pilot inputs. The rudder area was slightly reduced (2.5%) when the powered elevator was installed. The rudder pedal force and gearing are shown in figure 10. Maximum rudder deflection (single panel) is +25".

The lateral control system is completely new. Three separate surfaces are used t o produce the required rolling moments: ailerons with boundary-layer control, spoilers in front of the ailerons, and augmentor chokes. Figure 1 1 shows the positjon and function of each element. The ailerons are mechanically programmed to droop as a function of the flap deflection, as shown in figure 12(a).

Full droop is 35" and is reached at flap deflection of about 70". The maximum differential aileron deflection from the droop position is +17" with flaps up. As flaps are extended, the control wheel to aileron gearing is changed slightly (fig. 12(b)) and + 18.5" of differential aileron is available at 65" flaps with SAS on. Blowing boundary-layer control is used on the aileron to increase the effective- ness of both the ailerons and spoilers. The augmentor chokes are designed to control the lift of the augmented jet flap system by changing the exit area of the augmentor. Full lateral control reduces theexitarea as a function of flapdeflection t o amaximumof 55% onthe down-going wing (fig. 12(c)). The chokes function much the same as spoilers on a conventional airplane. Although there are augmentor chokes in each of the four sections of the flap, only the chokes in the outboard section of each wing are used for lateral control. All four chokes may be activated on the ground after landing for lift dump.

The choke and spoiler lateral control surfaces are programmed to give nearly linear effective- ness with control wheel deflection (fig. 12(d)).Thespoileroperatesfromnearneutral wheel positionandtheaugmentorchoke is phased in at about 11" (flapdeflection65")control wheel deflection. The spoilers are fully deflected at about 45" wheel deflection.

Lateral control wheel forces are low (fig. 13) and are produced by a simple spring system. The lateral control surfaces are activated by a central dual hydraulic power actuator, located on the rear spar, which drives the ailerons through a cable system. The central lateral power actuator also drives spoilers andaugmentorchokecontrol valves through a second cable. The spoiler and choke actuators are powered by separate hydraulic systems.

Approximately midway through the flight-test program several modifications were made to the airplane. The SAS gearing was changed to increase the Dutch-roll damping, the lateral control feel spring was modified to improve the centering and force gradient, and the poweredelevator system was installed.Thepowered-elevatorsysteminstallationcausedanincrease in the airplane'sroll, pitch, and yaw moments of inertia and a decrease of 0.14 m2 (2.5% of total) in rudder area. At a gross weight of 178,000 N (40,000 lb) the moment of inertia about the yaw axis was increased about 5% while the inertia increase about the roll axis was about 1%.

Rolling moment is produced by ailerons, spoilers, and augmentor chokes. Figure 1 1 illustrates the location and deflection of these controls. Gearing of the pilots control wheel t o these surfaces is shown in figure 12.

Maximum control surface rates of deflection are as follows: Z 2OO0/sec Control wheel Ailerons 2 5O"/sec 2 12O0/sec Spoilers

Chokes > 35"/sec

With flaps deflected for landing, full lateral control from neutral position can be achieved in about 0.5 sec.

There is an increase in effective gearing of the control wheel t o surface deflection from 0 t o +3" of wheel deflection;theincrease is producedbyacircuit in the SASsystem,asshown in figure 14. With SAS turned off this feature is disabled. With SAS on, the gearing is double that of the SAS off case for the first +3" of wheel deflection. This system was incorporated t o reduce any lag in aircraftresponse at smallwheel deflections in the region of smallspoilerdeflectionwhere theyare relativelyineffective. As theflapsareextended,thecontrol wheel t o ailerongearing is increased somewhat. Figure 12(b) shows the changein total aileron deflection available at various 65" flap deflection, the total aileron deflection available is about 10% greater flap deflections. At than that with flaps up.

Stability Augmentation System Twoindependentstabilityaugmentationsystems,oneforthelateralandoneforthedirec- tional aircraft axis are provided. Blockdiagrams of the SAS systems are shown infigure 14. The SAS actuators arepositionedbyclosed-loopservosandsummed in serieswith the pilot's control system. The SAS electronics is single channel, and relies on limited rate and displacement authority for safety. The lateral rate limit is 50"/sec with the displacement authority limited t o +20" (27%) equivalent wheel displacement. In the first +3" of equivalent wheel deflection, the SAS to control gearing is double that beyond 3". The maximum directional SAS rate is 25"lsec of rudder travel, anddisplacement is limited to +5" (20%)offore-rudderdeflection.Afterthepresenttests were completed, the displacement limitwas increased t o +9" rudder-deflection.

The lateral SAS in the normal mode performs the following functions: 1 . Spiral stability augmentation, using yaw rate feedback to the lateral controls 2. Roll damping augmentation, using roll rate feedback to the lateral controls 3. Lateral control quickening, usingwheel position feed forward (+3" 6 , maximum) to the lateral controls.

The directional SASin the normal mode performs the following functions: 1. Turn coordination, using roll rate and roll attitude feedback to the rudder 2. Dutch-roll damping, using yaw rate and roll attitude feedback to the rudder.

The SAS gains (fig. 14) are programmed with flap position and are automatically switched off above 100 knots. About midway through the test program the gain was increased from a value of 0.4 to 0.8 to alleviate a snaking tendency at STOL approach speeds. Figure 14(c) is the schematic for the lateral control surface quickener, which doubles the lateral control gearing for thefirst 3" of control wheel travel to improve the control characteristics near the neutral control surface position.

Data Acquisition System An on-board data acquisition system gathers data on about 95 parameters measured during all ground and flight testing. A pulse code modulated (PCM) digital system recordsthedataon magnetictape.Recordedflight-testdataincludestabilityandcontrol,SAS signals, engine gerfor- mance, augmentor performance, and guidance information. Each parameter is sampled 100 times per second.

The PCM data are processed on a digital computer programmed with the desired equations for determiningsuchparameters asengine thrust,liftanddragcoefficients,flight-pathangle,and corrected airspeed.

For certain tests, the on-board data acquisition system also included a recording oscillograph, which is used for recordinghighfrequencydata.This was used forflutter,vibration,andloads testing.

FLIGHT-TESTPROCEDURE Theflight tests were initiatedfromMoffett Field andflown in test areasin the vicinity of Moffett Field and at the Crows Landing Naval Auxiliary Facility. The flights were made by project pilots from NASA, the Canadian Department of Transport, de Havilland Aircraft of Canada Ltd., and The Boeing Company.

The flight tests included the following general categories: calibration, aerodynamics, stability and control, performance, and operational and handling qualities.

Stability and Control Tests Thestabilityandcontroltestconfigurations were generally cruise (5.6" flaps) at 150 knots, landing approach (65" flaps) at about 65 knots with enane nozzles at 75", and takeoff (30" flaps) at about 80 knots airspeed. In general, power for level flight was used for the cruise and landing approach testing, while maximum continuous power (96% rpm) was used for the takeoff configura- tion tests. Simulated engine-out testingwas conducted by throttling one engine t o idle.

Stability and control tests were performed to determine the following: control system characteristics,controlpower,controlsensitivity,staticstability,dynamicstability,acceleration characteristics, and trim changes.

Longitudinal test maneuvers included the following: trim change with speed (with one and two engines); trim change with power and with nozzle deflection; elevator steps, reversals, and doublets; rapid pitch attitude changes; wind-up turns; engine power lever steps; engine nozzle lever steps; trim change with flap deflection; phugoid; trimchange in ground effect; and stall behavior.

Lateral-directional test maneuvers included the following, with SAS both on and off: lateral control steps and reversals; directional control steps and reversals; Dutch-roll; spiral stability; rapid bank angle changes; trim change with one engine operating; stalls; and steady sideslips.

Tests were also conductedtodeterminetheeffectsofpartialhydraulic failure on lateral control and on SAS operation. The lateral control augmentor chokes and the spoilers were deactivatedseparately to testtheireffectoncontrolpowerandsensitivity.Thedirectionaland lateral control channels of the SAS were deactivated separately, again t o determine their effect on stability and control.

Operational and Handling Qualities Operational and handling-qualities testing was conducted in conjunction with the other tests.

In addition, STOL landing, transition, takeoff, and waveoff operation, and ground effect testing were conducted as part of the evaluation of operationalandhandlingqualitiescharacteristics.

Simulated single-engine landing approaches and takeoffs were also conducted.

FLIGHT-TEST RESULTS AND DISCUSSION The results of the flight investigation of stability and control are discussed under two categories: ( 1 ) longitudinal stability and control, and (2) lateral-directional stability and control.

Longitudinal Stability and Control Longitudinalstabilityandcontrolcharacteristics of theairplanearesummarized in tables through 7. Stability and control derivatives for the landing approach configuration are tabulated in table 6. Thesestabilityandcontrol derivativesare fromreference 13 and were obtainedfrom flight-test data recordings using an equation error parameter identification method and from prediction.

After the first 78 hr of flight testing, a poweredelevator systemwas installed in the airplane to replace the spring-tab system. In addition to the direct effects of the powered elevator on stability and control, the modification changed the airplane's center of gravity and moments of inertia. The pitching moment of inertia was increased about 14% at 178,000 N (40,000 lb) gross weight and the center of gravity was moved aft 0.3%. The variation of center of gravity with gross weight is shown in figure 15.

LongitudinaZ controZ- Beforemodificationof the elevatorcontrolsystemthepilotsfound thatthelongitudinalcontroldynamicfeelcharacteristicswereunsatisfactory,andthemaximum control from trim that could be obtained at STOL airspeeds with reasonable forces was only about +lo" to - 17"elevatordeflection,althoughstatically,fullelevatortravel was +15" to -25". For subsequent testing, the elevator control system was modified to a fully powered system (ref. 12).

Controlcharacteristicsof the aircraftwith the spring-tabsystemarepresented in references 10 and 1 1 . The data presented in the followingdiscussionis for the poweredelevatorconfiguration except where otherwise noted.

The longitudinal pitch control characteristics of the aircraft are summarized in the following table. (The control power is for - 24" elevator deflection.)

LONGITUDINAL PITCH CONTROLCHARACTERISTICS Airspeed, knots 60 65 75 90 0.47 X . 6 Control power, rad/secZ 0.41 X . 6 Sensitivity, rad/sec2/cm .027 .032 .043 .062 rad/secz/in. .070 .082 .lo9 .157 1 0 . 1 11.1 14.0 Force gradient, N/cm 9.5 lb/in. 5.4 5.8 6.3 8.0 Breakout force: N 25.8 25.8 26.7 27.6 5.8 lb 5.8 6.0 6.2

I

aIncludes system friction of 17.8 N (4 Ib).

Theangularacceleration fordeterminingcontrolpowerand sensitivity was obtainedfrom longitudinal control reversal maneuvers of various amplitudes. A typical control reversal at 67 knots is illustrated in figure 16. The angular acceleration for a specific change in elevator position from trim was obtained by the slope of the pitch rate at the time pitch rate is zero, where the damping term is zero. Corrections for changes in angle of attack are made where appropriate. The elevator control effectiveness is shown in figure 17. Figure 17(a) presents the data in terms of the variation of the ratio of angular acceleration to the free-stream dynamic pressure ((ilif) with the incremental elevatordeflectionfromtrim.The variation ofangularaccelerationwithincrementalelevator deflection,computedfromthedata in figure 17(a)at60,65,75,and 90 knots is shown in figure 17(b). The data show that for the landing approach speeds of 60 to 65 knots the maximum control power is between 0.41 and 0.47 rad/sec2. At 75 knots and above the angular acceleration exceeds 0.6 rad/sec2. The pilots felt that angular accelerations greater than 0.5 rad/sec2 could not be used effectively except in emergencies.

... . . .

The gearing of the longitudinal control was 1.7" elevator per degree of column deflection for small angles at 60 knots. Some small nonlinearity is evident near zero degrees and at high deflection angles, andthere is a change in gearingwithairspeed.Detailsof the systemarepresented in reference12.

The length of the column is 0.83 m (32.5 in.). The force characteristics are shown in figures 6 and 7 and in the table of longitudinal characteristics (table 2). The pilots considered the control feel characteristics of thepoweredelevatorsatisfactoryand to besuperior to the spring-tab control system.

of the powered elevator and the original spring-tab A comparison of the control effectiveness elevator,figure 18, showsthepoweredelevator effectiveness is about 14% less than that of the spring-tab elevator. This reduction results from th.e increase in pitching moment of inertia due to the concentration of the powered-elevator system components in the tail of the aircraft and the additionof ballast in thenoseoftheaircraftfor balance. The aerodynamicfeaturesofthe horizontal tail and elevator were not significantly changed.

Thepilotsconsideredthelongitudinalpitchcontrolcharacteristicssatisfactoryfor all STOL operation. Hysteresis in the engine power control system produced unsatisfactory thrust control in STOL landing approaches. Elevator deflections as high as -18" (about 15" from trim) were noted in some takeoff rotations and landing flares. The control power was adequate to recover from all stalls conducted. However, at very slowstallspeeds,41-43 knots, the pilots used full forwardcontrol which gave a slow recovery. The project pilots, therefore, considered 45 knots the minimum flight speed for the airplane to ensure adequate control.

Longitudinal static stability- The installation of the powered-elevator system in the airplane hasremoved the undesirable stick-free characteristics of the spring-tab elevator system. Stick-free characteristics.of the airplanewiththespring-tabelevatorarepresented in reference 11. The stick-fixed stability characteristics will be discussed in the following paragraphs.

The static longitudinal stability characteristics of the airplane were obtained by flying steady trim conditions from flap placard speeds down to and including the stall and measuring the elevator angle required for trim. Tests were conducted over a range of flap, nozzle, and power settings. The center of gravity varied between 30.0% and 30.9% of mean aerodynamic chord. Test weights varied from 178,000 to 205,000 N (40,000 to 46,000 lb). Figures 19 through 2 1 present elevator-to-trim data for a broadrange of conditions, including simulated singleengine operation.

Inthelandingconfigurationwith flaps at 65" (fig. 19),nozzleseitheraft or down, with of between 90 and 99% rpm, trim is achieved for all conditions symmetrical power on both engines testedwithelevatordeflectionswithin 6" of neutralat all angles of attack up to thestall.The airplane is staticallystable at anglesof attack less than 10" for all conditions of constant engine power or constant nozzle deflection tested and for all angles of attack at engine rpm below about 93%. Above 93% rpm and at high angles of attack, the stability can either be positive, neutral, or negative depending on the combination of nozzle angle and enginepower. In general,increasing engine power has a destabilizing effect and increasing nozzle deflection can either be stabilizing or destabilizing.Neutral stability is evident,however, at angles of attack between 10" and 15" with 95% rpm and nozzle angle equalto or less than 67"; for the same conditions, positive stability exists with the nozzle deflection equal to or greater than 88". The airplane is unstable with rpm greater than 95% and nozzle deflection eq.ual to or less than 76" at angles of attack between 15" and the stall angle. A strong nose-down pitching moment exists at the stall for most configurations.

In the simulated engine-out configuration with landing flaps (65"), positive stability is evident (fig. 19(d)) for all conditions tested except for a region of neutral stability in the angle of attack range near 15" with the highest engine rpm tested (99.6%) with a nozzleangle of 72". With a similar rpm (96.5%) there is reduced but positive stability in the nozzle deflection (73") at lower engine region between 10" and 15"angleof attack.Thereasonfortheapparentnonlinearity is not understood.

The elevator to trim data for the takeoff flap configuration (30") with nozzles up is presented in figure 20 for two engines operating and .for the simulatedone-engine operating conditions. In both cases the airplane is static'ally stable throughout the angle of attack range tested except for the highest engine power condition (two engines operating). With 99% rpm on both engines (fig. 20(a)) there is a neutral stability region near 15" angle of attack with positive stability evident at angles of attackgreaterthan 20" and less than 12".In the region of the stall there is a strong nosedown pitching moment for conditions tested with less than 94% engine rpm. Some nose-down moment near the stall is also indicated at higher engine power. The region of neutral stability at intermediate angles of attack was not evident for simulated one engine operation (fig. 20(b)). At either constant airspeed or angle of attack, increasing engine power produces a nose-up pitching moment.

With the flaps up (climb and cruise configurations), the airplane exhibits positive static rpm (approximately stabilityasshownin figure 21. Figure21(a)showselevator-to-trim for 90% cruise power), and for 97% r p m , which is 1 % higher than maximum continuous power. For both conditions of engine power,thestabilitybecomesquite high at angles of attack above 20" with large aftcolumndeflectionsrequirednearthestall.Thesimulated single engine condition also is stable (fig. 2 1 (b)) for the97%/idle rpm case tested.

Longitudinal dynamic stability- Thedynamiclongitudinalshortperiodstabilitycharacteris- ticswere approximated by analyzingthe responseof the aircraft to control pulses andsteps.

Aircraftshort-periodandphugoidcharacteristicsaresummarized in table 5 for several configura- tions. Figure 16 is a typical time history of response to an elevator step at 66 knots airspeed. The data illustrate the low short-period dynamic stability of the airplane at this speed. The response to a stepelevatorappearsto be almost first orderwithatimeconstantofabout 0.5 sec;pitchrate becomes almost constant after about 1 sec. Stability values shown in table 5 were computed by the method of reference 13 and show the short-period mode to be of the order of 10 to 16 sec and a dampingratio of about 0.90. The pilots statethatthere seems to be very little if anystability.

Figure 22 is a time history of a phugoid oscillqtion from a trimmed condition of 65 knots with the nozzles deflected 77". The period of the phugoid mode is over 30 sec with a damping ratio of less than 0.10.

Thepilots were not satisfiedwith theshort-perioddynamiccharacteristics of theaircraft because of the high pilotworkloadrequired to maintainadesiredpitchattitudeand t o control airspeed onthelandingapproachatairspeedsof 60 to 70 knots. Approachesandlandingsunder visual flight rules (VFR) were performed in windsas high as 30 knots withgusts.Flight in gusty conditions resulted in only small low-frequency disturbances in attitude because of the low dynamic stability. The pilot workload increased, however, during the approach because of flight-path disturbances.

Flight-path controZ- Theflightpath of the airplaneinafixedflapconfigurationmaybe controlled by either modulation of engine power (rpm) or engine nozzle deflection. Normally, the engine nozzles are used for flight-path control when flying in the landing approach configuration; however, they can be used effectively for longitudinal acceleration control in other flap configura- tions. The variationofflight-pathangle with airspeed forthelandingapproachconfiguration is presented in figure 23 for a typical airplane gross weight of .178,000 N (40,000 lb). Reference 11 presents the flight-path/airspeed relationship for the takeoff and cruise configurations which show characteristicssimilar to those of conventional airplanes.Figure23(a)shows the change in flight pathand angle of attackasafunctionof airspeed andengine rpm forthelandingapproach 90”. Thenozzledeflection used fortypicalSTOL configuration (65O flaps)withnozzlesset at landing approaches at 65 knots on a -7.5” flight path is nominally about 75”. The data show that the aircraft exhibits “back-side” characteristics at 93% engine rpm required for a -7.5” approach rpm reduces the onset speed at which pathat airspeeds less than72knots. Higher engine “backside” conditions exist and the converse is true for decreasing engine rpm. It is also evident that for a constant airspeed approach in the 65 knot speed region, the flight path may be changed by rpm adjustment with a small change in pitch attitude opposite to the direction of the flight-path angle change. If power is increased t o decrease rate of sink, a nose-down pitch attitude change is required to maintain airspeed. The trim angle of attack change is also of opposite sign but of similar magnitude to the flight-path angle change. These characteristics are typical of propulsion lift aircraft where a substantial amount of liftis produced by engine thrust.

Theflight-path/velocityrelationship as a function of engine thrust vector angleis shown .in figure 23(b) for the fixedrpm (94%) landing approach Configuration. Large changes of flight path can be made by a change of nozzle position at constant airspeed withessentially no change in angle of attack. The flight-path angle change is therefore proportional to pitchangle change when nozzles aremodulated to maintainairspeed. It is alsoevident that“backside”conditionsexistforany nozzle deflection at airspeeds less than about 72 knots. The nozzle control can be used much the same as a throttle in a-conventional jet aircraft if pitch attitude is coordinated.

In practice, boththethrottlesandnozzlesaremodulated in thecontrol of flight path, airspeed, and/or angle of attack on a landing approach. In general, the nozzlesareused to make large flight-path corrections, and pitch attitude is used for airspeed control. Throttles are used t o make small corrections to glide path and/or airspeed with pitch attitude used to control angleof attack. A timehistory of atypicallandingapproach is shown in figure 24 which shows the last 60 sec of flight before touchdown when the nozzles were used to decelerate from 90 knots to the approach speed of about 66 knots. The nozzles are moved only once after decelerating t o near the approach speed while the rprn is adjusted several times for flight-path control. The approach angle of attack is nominally about 4” with some random excursions which may be attributed to reported turbulence. Little elevator is used throughout the approach until the final landing flare where a peak deflection of about 14” from trim is used.

LongitudinaZ maneuvering characteristics- This section covers the aircraft’s vertical and longitudinal acceleration responses to control inputs used for flight path, flare, andairspeed control.

Four controls are available to the pilotinthispropulsiveliftaircraft t o achieve these responses: (1) changes of angle of attack with elevator control, (2) changes in engine thrust with the throttles, (3) changes in hot thrust vectorangle with nozzle control levers, and (4) change in lift and drag with flap deflection. Each of these controls affects the lift, drag, and pitching moment of the aircraft.

The aircraft exhibits stable maneuvering characteristicsat constant airspeed, thrust, andnozzle angle fortheconditionstested.Figure25comparesmeasuredmaneuveringdatawithpredicted values fromreference 14 in terms of angle of attack-per-g, elevator position-per-g, andstick force-pes-g variation with airspeed obtained from wind-up turns and pitch attitude steps. Control force data are presented only for the poweredelevator system.

Load factor per unit angle of attack (nZa) becomes quite low asairspeed is reduced and is slightly less than predicted. The test values of nza, which were obtained from pitch attitude step maneuvers, arefor conditions where Aor, is less than about 4.5". In the maneuvers in which AcyF was 5" or greater, the airspeeddeviationfromtrim and pilot technique havea large influence on the measured value of Anz. The elevator-per-gis greaterthanpredicted.Thegreaterthanpredicted elevator-per-g may be accounted for by the lower than predicted load factor per unit angle of attack change. Typical wind-up-turn test data are presented in figure 26 for the takeoff configuration.

Figure 27 shows peak normal acceleration achieved during step changes in pitch attitude with flaps at 65". A comparison of the test data is made with computed values based on the change of CL with angle of attack at a constant CJ of 0.4. The computed values are based on a trim angle of attack of 5" and a constant airspeed; although test values of trim angle of attack varied up to lo", the data are applicable because the lift characteristics are quite linear for angles of attack between 5" and 10". At 63 knots airspeed (fig. 27(a)), good agreement exists between test values and com- puted values of An, for conditions of AQF less than about 5". At 63 knots, for test conditions of AaF greater than 5", the measured values of Anz tend to be lower than computed due t o a reduc- tion in airspeed in the maneuver. At 68 knots airspeed (fig. 27(b)), test values of An, tend to be greater than predicted.

Figure 28(a) shows time histories of 5" and 10" pitch attitude changes with 65" flap deflection at a trim airspeed of 63 knots. For the 10" step, note that as pitch attitude is increased, the airspeed decreases such that the peak value of vertical acceleration does not occur at the peak value of angle ofattack. Also,analysis ofthetestdatashowthatpilottechnique,phasing,andamplitudeof control inputs has an influence on the peak An, value, particularly for the larger (>5") test values of AaP These effects account for the scatter in An, with ACXF shown in figure 27. With attitude changes of .So, acceleration follows angle of attack more closely; airspeed changes are much smaller and do not affect the peak load factor. Figure 28(b) presents similar time history datafor nose-down attitude changes.

Figure 29(a) shows aircraft response to thrust vector nozzle rotation downward from the aft position, with the pilot controlling elevator t o hold angle of attack nearly constant. The effect is somewhat similar to a decrease of thrust in a conventional aircraft because the main effect on the flight characteristics of vectoring thrust is to change the thrust minus drag relationship. When the nozzle is moved from full aft t o 90" deflection on the 7.5" landing approach path, the liftchange is only about 35,000 N which is about 20% of the weight of the airplane. As thethrustvector is rotated from aft to verticalposition the airspeeddecreasesquicklywithlittlechange in angleof attack, then increases as the aircraft pitches down and rate of sink increases, exciting the phugoid.

The new trim speed is lower than initial trim because of the added lift due to thrust deflection. If pitch attitude were held constant, the flight path would steepen, angle of attack would increase, and speed would decrease. An initial increase in positive vertical acceleration would accompany a more rapid thrust (nozzle down) deflection. The effect of rotating nozzles from the near vertical to aft position is shown in figure29(b). Thetimehistoryshows that the pilotinitially puts in forward column to counter the nose-up trim change. The aircraft sinks momentarily as the lifting force due to thrust isremoved (about -0.1 g),speedincreasesrapidly,and the aircraft climbsas the pilot pitches the aircraft up. With the exception of the initial increased sink rate due to the lift changes, the response is similar to a thrust increase in a conventional aircraft. Figure 29(c) illustrates similar effectsforasmaller' change in nozzleposition.Thrustvectoring was aneffectivemeansfor deceleration to approach speed and for flight-path control during glide-slope interception. Once on a -7.5" approach path at 60 t o 65 knots, modulation of the nozzles in the range 70" t o 90" provided a very effective airspeed or flight-path control which could take the place of conventional throttles.

The initial tendency for the aircraft to sink when vectoring the thrust aft, however, restricted the use of the nozzlesasa controlinthefinalphase-oftheflare close to theground.Thepilots preferred to revert to throttleandelevatorcontrolforthetouchdown.A special technique is required in low altitude waveoffs to avoid sinking when the thrust is vectored aft; increasing angle of attack as nozzles are moved aft effectively alleviates a problem with theloss in lift.

Aircraft response t o changes in throttle settingis conventional in the cruise configuration. With flaps down,theresponse is unconventional, particularly in landing configuration with highly deflected flaps and nozzles down. Figures 30(a) and 30(b) illustrate response to a step increase and a step decrease in power, respectively, in the landing configuration. In the time histories shown in figure 30, the pilot is controlling with the elevator to maintain constant pitch attitude. Increasing to 83" produces a lifting force, instead of an axial force, thus causing thrust with nozzles deflected the aircraft to heave upward. Figure 30(a) shows a peak vertical acceleration of 0.10 g for a step r p m at2060m (7000 ft)altitudeandno changeinlongitudinal fromabout92.5%to98.5% acceleration.Flightpathshallowsinitially, but with a near constant angle of attack the airspeed decreases tending t o wash out the change in flight-path angle.

Similarly, the short-term effect of reducing power (fig. 30(b)) with attitude held constant is a decrease in load factor and a steepened flight-path angle, butas speed increases the flight path tends t o shallow. With the use ofthrust as theprimaryflight-pathcontrol,this adversespeed-path couplingmadeitnecessaryforpilots to continuallycontrolbothattitudeandthrust to achieve desired corrections in flight path and speed. Where the required flight-path corrections were large, pilotspreferred to revert to acombination ofnozzleandthrustcontrol to achieve the desired response. The direct pitching moment produced as a result of engine throttle or nozzle changes is small because of the proximity of the thrust reaction to the airplane center ofgravity. The airplane response to changes of flap deflection is docile because the rate of change of flap deflection is low (4"/sec).Theelevatorrequired to trimasafunctionofairspeedforfourflapconfigurationsis shown in figure 31 for a gross weight condition of approximately 197,000 N (44,300 lb) at about 95% r p m . Shown also in figure 3 1 is the variation of elevator required at constant angles of attack as flaps and airspeed are varied. The trim change is nose-up at constant airspeed as flaps are lowered requiring forward column deflection for trim. A change of about 5" (airplane nose-down) elevator deflection is requiredforatransitionatconstant 5" angle ofattackfromcruiseconfiguration (148knots,flapsup) to thelandingapproachconfiguration(63knots, 65"flapdeflection, 88" nozzledeflection).Typically,thetransition is madewith sizeablevariationsinengine rpm and nozzle deflection and small variations of angle of attack.

A time history of a transition from about 30" flaps t o landing flaps is shown in figure 32. The datashowthatalthoughflaps,nozzles, engine rpm, and airspeedareallchanging, only small elevator angle changes (about 6") are required. About 7 sec are required t o change the flap deflection the first 20". The last 10" change of flap deflection takes about 14sec.

Longitudinalresponse to asimulatedenginefailure- Theairplane response t o asimulated enginefailure is presented in figure 33 for the takeoff and landing configurations. In the takeoff configuration (fig. 33(a)) the starboard engine is throttled back just after lift-off, withflaps at 24.8" deflection and about 90 knots airspeed. The gross weight was 210,000 N (47,000 Ib). As the engine rprn spins down to idle, the rate of climb decreases from 7.5 m/sec to near zero and angle of attack increases from 6" to 10". Only about 2" of up elevator is required t o maintain trim. During the entire maneuver the airspeed is held within 5 knots of 90 knots, which was the airspeed at the time the throttle was moved to idle.

Thelandingapproachsimulatedenginefailuretimehistory is showninfigure33(b)fora 165,000 N (37,000 lb) gross weight condition. The starboard engine is retarded from 95% rprn t o idle with flaps at 65", nozzles at 65", and airspeed about 70 knots. As the engine rprn spins down, the nozzles are rotated aft and airspeed is increased to 80 knots, at which time the flap retraction is started.The angle of attack is heldnear 2" untilthestartofflapretraction where the angle of attack is increased to about 6". The elevator required for trim is 2" to 3" until the beginning of flap retraction. At 30" flap deflection (92 knots) the required elevator is -2". A positive rate of climb is achieved at about 92 knots airspeed with 99% rprn on the port engine at a density altitude of 504 m (1 654 ft). About 150 m of altitude is lost in the recovery. The pilots considered the trim change andpilotworkloadwithasimulatedenginefailurelittledifferentfromconventionalaircraft;the performance, however, was quite different. In the simulated engine failure, the thrust change to idle is slow (20 sec) due to the deceleration schedule of the engine. A rapid deceleration of the engine in anactualenginefailurewouldincreasethepilotworkloadas the trim change and lift loss would occurmorerapidly.Recognitionof enginefailurewould bedifferentfromthatexperienced in conventional airplanes as little or no side force is noted by the pilots.

Longitudinalcharacteristicsin the stall- Thecharacteristics at the stallaresummarized in table 7 for various takeoff, cruise, and landing configurations. Time histories of approach to stalls arepresented in figure 34. Ingeneral,thestall is characterizedbyamildreduction in vertical acceleration preceded by buffeting as the stall is approached. Most configurations exhibit a strong nose-down pitching moment near the stall or just beyond the stall angle of attack. Recorded angles of attack at the stall varied between 17" and 30", depending on airplane configuration. The lowest angle of attack recorded at the stall (17') was experienced in the landing configuration with one engine at idle. The highest stall angle of attack recorded was in the cruise configuration. The only adverse longitudinal characteristic observed in the stall was for the landing approach configuration with high engine power, in which case a slight pitch-up tendency was noted. In the case shown in figure 34(a), the pilot used fullforwardcolumn to recoverfrom the stall and expressedconcern because of the sluggishness of the angular response in pitch. For this condition the airspeed was about43knots, nozzleswere73",engine rprn was 99.4%,and thetrimelevatorat 5" to 6" nose-downas the stall was approached; the airplane gross weight was 189,000 N (42,500 Ib).In most other configurations the airplane had a tendency to pitch down as stall was approachedas shown in figs. 34(b), 34(c), and 34(d). It should be noted that there is considerable uncertainty as to airspeed near the stall. The airspeed calibration (ref. 1 1 ) indicates inaccuracy at large values of CL.

With both engines operating at 96.7% rprn in the landing approach configuration (65" flaps) with engine nozzles at 74", the stall angle of attack is 27" (fig. 34(b)) and airspeed at the stall is 47knots.Theairplane gross weight forthisstall was 191,000 N (42,900 Ib). As the stall is approached, the airspeed remains constant at 4 7 t o 48 knots and the elevator required for trim is . I essentially constant from 21" angle of attack to the stall. Recovery from this stall was made using 13" nose-downelevator,which was initiatedat 30" angle ofattack (CL = 5.5)where CL has 6.3 at the stall.

decreased from a maximum of Themaximumenginepoweratwhichthestall was recordedwithtakeoffflaps(30") was 93.5% rpm. Athigher engine power,thepitchattitude became very high (>30") andthepilots preferred not to conduct full stalls. A time history of the approach t o stall is shown in figure 34(c) for a gross weight of 192,000 N (43,200 Ib) with 99.6% rpm. In this stall approach, the maximum CL of 4.5 isachieved at 25" angle of attack, 29.5" pitchattitude,and52knotsairspeed.Stall recovery was initiated at 25" angle of attack. About 3" nose-upelevator is required for trim at maximum lift coefficient and 1 1" nose-down elevator increment is used t o effect the recovery. The extrapolated stall speed that would be predicted at sea level with NH at takeoff rpm (CJ = 0.92) and for aweightof 200,200 N (45,000 Ib) wouldbe 50 knots and require about-2" ofelevator deflection; pitch attitude would be about 27". The maximum test value of CJ at the stall was 0.62 with NH at takeoff rpm and flap deflection at 30".

In the cruise configuration (flaps up) with 90.4% engine rprn, a maximum lift coefficient of 2.9 is reached at 28" angle of attack. Figure 34(d) is a time history of the approach t o stall for this condition.Themaximum liftcoefficient of 3.1 was reached at 69 knotswith an airplane gross weightof 197,500 N (44,400 lb). An increasing amount of aftcolumndeflection is requiredas maximum lift coefficient is approached, with - 14" elevator required for trim at the stall. Beyond maximum lift coefficient at 30" angle of attack the lift coefficient is reduced t o 2.7 at which point recovery is initiated using about 7" nose-down elevator.

Ingeneral,thepilotsfoundnoproblemwiththecontrolcharacteristicsatthestalland conventional recovery techniques were applicable. There was some lateral and directional unsteadi- ness at angles of attack near the stall which required constant attention of the pilot. The control task was not considered difficult, but the work load was high.

Ground effect- In the landing configuration, measurements of ground effect show a marked decrease in drag, a nose-down pitching moment, and a slight increase in lift as the airplane descends to ground level. Figure 35 shows theeffect of groundproximityonthelift,drag,andpitching momentcoefficientobtainedfrom several landings in which theairplanedescended slowly to touchdown.Thediscretedatapointsshown werederived fromtestdata (assumedquasi-steady- state) collected with flaps 65", C L ~ = 2.3 and 3.0, and engine nozzles deflected between 50" and 80". The data points for C L ~ = 3.0 are typical of the STOL landing conditions; however,in practice the flight path is much steeper. Separate flight conditions were used for determining lift, drag, and pitching-moment ground effects. The engine power varied between 93% and 96% rpm, depending on flight conditionsand gross weight.Figure 35 alsopresentsfairings (CL = 2.65)based on a A least-squares fit to the entire data for four runs using a regression parameter identification technique (ref. 15). When the airplane is flying at ground level the drag is about 70% and lift 105% of the basic values (CL = 2.65 and C o A = 0.45). The nose-down pitching moment at ground level for the same condition IS equivalent toabout 8" elevatordeflection (ACm = -0.3).Thepitchingmoment in-ground effect was not considered excessive bythe pilot.Reference 15further discusses the groundeffectandshowsa sizeablechange in thegroundeffect (CmG and C L ~ ) withangle of attack. These data indicate ACmG/ACX = - 1.l/rad and A C L ~ / A C X = - 1.O/rad between - 1" and 6" angle of attack with the airplane flying at ground level (6f= 65"). There was no measurable change in C o G with angle of attack for the test conditions.

Lateral-Directional Stability and Control Lateral-directional stability andcontrolcharacteristicsoftheairplanearesummarized in 8 through 12. The stability derivatives for the STOL landing approach condition at 62 knots tables airspeed are given in table 8. These stability and control derivatives are from flight tests with the SAS turnedoff,which were performedafterthepoweredelevator was installed. A parameter identificationmethodsuch as thatreported in reference 1 3 was used t o obtain the stability derivatives.

Lateral controZ- The lateral control effectiveness is a function of several variables in addition to freestream dynamic pressure (airspeed) and the three separate control surface deflections. The rolling moment due to aileron deflection varies with the momentum coefficient of the boundary- layer control (BLC)air, and, therefore, varies with engine fan thrust (engine speed, altitude, and temperature), and aileron droop angle, which is a function of flap angle. The rolling moment due to spoiler deflection varies, in turn, with the effectiveness of the drooped ailerons. The rolling moment duetotheaugmentorchoke varies alsowithenginefan thrustandflap angle. In addition,the angular acceleration for a specific total rolling moment dependson the roll moment of inertia which changes significantly with distribution of the fuel load in the wing tanks. Also, under singleengine conditions,onlyoneaileron has BLC because ofthe crossductingarrangement.It is difficult, therefore, to define the lateral control effectiveness for all conditions. Data will be presented that are for configurations that are similar to those used in STOL takeoff, landing, andcruise.

Lateralcontroleffectiveness was measuredbyconductinglateralcontrol reversals with the airplane in variousconfigurations. A force was applied to the control wheel in one direction and then rapidly reversed and held in the opposite direction with rudder pedals neutral. Rolling angular acceleration was measured as roll rate passed through zero. A typical reversal at 65" flaps with SAS off is presentedinfigure36(a)andwith SAS on in figure 36(b).Themeasured variation of maximum lateral control power with airspeed for the takeoff, cruise, and landing configurations is presented in figure 37 alongwithpredictions.Thepredicted values for the flighttestpointsare shown for 65" and 30" flap configurations (figs. 37(a) and 37(b)) and are within 0.01 rad/sec2 of the test values except the 92% rpm, 30" flap case shown. The flight-test measured control power at 92% rpm with 30" flaps is about0.04rad/sec2greaterthanpredicted. Since there is excellent agreement between test and predicted control power, the predicted control power curves shown for sea level standard day conditions are representative of actual airplane performance. An exception is the takeoff flap (30") condition at reduced engine power (90% rpm) where the predicted control power appears conservative. The predicted curve shown in figure 37(c) for lateral control power for the cruise Configuration ( 6 f = 5.6") is based onthe average of the measured values ofthe rolling-moment coefficient computed for a rolling moment of inertia of 474,500 kg-m2 (350,000 slug-ft').

In the approach configuration at 65 knots airspeed (65" flaps, 93% rpm) the maximum rolling acceleration at 178,000 N (40,000 lb) gross weight is 0.63 rad/sec2. At 200,000 N (45,000 lb) the maximum rolling acceleration is 0.50 rad/sec2 at 65 knots and 92% rpm. The lower value for the higher gross weight reflects the increase in roll moment of inertia.

In the takeoff configuration (30" flaps) at 90 knots airspeed the maximum rolling acceleration available with 100% rpm is 0.7 1 rad/sec at 200,000 N (45,000 lb) and 0.86 rad/sec2 at 178,000 N (40,000 lb) gross weight.

The maximum rolling acceleration measured with flaps up was 1.33 rad/sec2 at 160 knots and 200,000 N (45,000 lb) gross weight. At 120 knotsairspeed a value of 0.85 rad/sec2 was measured at 196,000 N (44,000 lb) gross weight.

Measured values of rolling acceleration as a function of control wheel deflection are shown in figure 38 for the 67" flapconfiguration at 69 knots. This figureincludesdataforthecomplete lateralcontrolsystemoperatingandforconditions where the chokes and spoilers were disabled separately to show the influence of each. The faired curves of figure 38(a) were used to construct the curves of figure 38(b),whichshowsthe rollingaccelerationproducedbyeachof the three lateral control surfaces. This figure shows that each control surface contributes about one-third of the total rolling moment for wheel deflections greater than40". At wheel deflections less than 40°, the contribution of each control surface reflects the gearing schedule shown in figure 12. The effect of engine power and SAS on lateral control is shown in figure 39 for all control surfaces operating at 69 knots airspeed. At 89% rpm with 60" of wheel deflection, the rolling acceleration is about 0.58 rad/sec2; at 93% to 95% rpm i t is about 5% greater. In the control wheel deflection region up to about 30", the SAS-on configuration should exhibit a greater rolling acceleration than does the SAS-off condition because of the "quickener" (fig. 14(c)); however, the differenceis not discernible in the data. In figure 36(b) it is evident that the SAS input to surface deflection at zero roll rate (the acceleration measuring point) is leading the control wheel deflection. The SAS-on roll to account for this by the inclusion of an equivalent wheel deflection acceleration data is corrected equal to the amount of the SAS input. With allsurfacesoperating (SAS off), the lateral control 2 rad/sec2 /degof wheel deflection in the region of less than 20" sensitivity at 69 knotsis about 0.01 wheel deflection.Figure 40 presentsrollaccelerationdatameasured at aspeed of 62 knots, and figure 41 summarizes 65" flap data in terms of the ratio of roll acceleration to dynamic pressure G/@. Available acceleration decreases as speed decreases, approximately in proportion to dynamic pressure, to about 0.5 rad/secZ at 60 knots. The roll control sensitivity is considered satisfactory by the pilots.

Figure 4 2 presents roll acceleration data for landing flaps; the data were obtained at 69 knots airspeed and a gross weight of about 200,000 N (45,000 lb). Maximum acceleration is reduced t o about 0.53 rad/sec2 at thisweight because of the increased inertia.

Figure 43 shows roll acceleration available in a simulated engine-out configuration with flaps at 66". Lateral controls are still very effective with a maximum of 0.59 rad/sec2 measured at 72 knots, 176,000 N (39,500 lb) gross weight, with the left engine at 61% rpm (idle) and the right engine at 99.7% rpm. For this configuration, about 10" right wheel deflection is required for trim and the engine nozzle deflection for maximum rolling acceleration to the right is less than to the left. The this case was 75", which is a more severe out-of-trim condition laterally than with reduced nozzle deflection. When the nozzles are rotated aft, the asymmetric moment is changed from roll to yaw.

With nozzlesundeflectedthelateraltrimrequired is about half that of the 75"nozzledeflection condition.

Maximum roll rates achieved in the wheel reversal maneuvers at flaps nominally 65" are given in figure 44 for SAS-on and SAS-off conditionsatabout173,000 N (39,000lb) gross weight.

Maximum rate available, SASon, is about 23"lsec at 69 knots.

Figure 45 presentsrollangularacceleration data for the 33" and 5.6" flapconfiguration at gross weights of 196,000to 200,000 N (44,000to45,000 lb). With 33" flaps, themaximum acceleration at 78 knots is 0.53 rad/sec2. These data are for an engine thrust setting of approxi- mately level flight; the control power with takeoff thrust is 0.59 rad/sec2 at the same airspeed and gross weight. With flaps u p (fig. 45), the measured maximum rolling accelerationis 1.25 rad/sec2 at 166 knots (Cl= 0.074). This high value is due in part to the lower surface of the chokes acting as ailerons. The choke deflection is only 10.4" with flaps up (fig. 12). Figure 46 presents additional data for approximately 30" flaps showing the effect of choke and spoiler on control power. At 60" wheel deflection the spoilers produce about 35%, the chokes about 25%, and the ailerons 40% of the total rolling moment. The control surface to wheel gearing is shown in figure 12.

In figures 36(a) and 36(b) it can be seen that with SAS on or off at STOL airspeeds, the yaw acceleration due to lateral control is very small. However, some adverse sideslip is evident; it will be discussed later.

Directional control- Directional control power was evaluated by conducting rudder reversals withthecontrol wheel at neutral.Maximumyawacceleration was measuredasyawrate passed through zero. Themeasuredyawacceleration is corrected toaccountforthe yawingmoment produced by the sideslipanglepresent at the time of measurement. A typical directional control reversal with SAS off is shown in figure 47, and with SAS on in figure 48. Roll acceleration due t o directional control with SAS on is small (fig. 48).

The rudder power measurements shown in figure 49 for the landing configuration are about 70%ofthosepredicted for one-half maximumrudderdeflection.At 14" ofrudderdeflection, 0.1 5 rad/sec2 yawing acceleration was measured at 72.5 knots airspeed (flaps 65"). The installation ofthepoweredelevatorreducedtherudderareaby0.14m2(2.5%)andincreasedthe yawing moment of inertia about 5% at 178,000 N (40,000 lb) gross weight. It is obvious that, aside from the inertia and area change, the empennage modification reduced the rudder power. After modifica- tion, the measured rudder power is about 70% of predicted, while it was 90% to 95% of predicted before the modification.

Results of rudder reversal testswith 67" and 33" flaps,whichwereconductedprior to the empennage modification, are compared with predicted results and are presented in figure 50. At 67" flapsand at anairspeed of 70 knots, ayawacceleration of 0.22rad/sec2 was measuredfor 17" (68%) of rudder deflection. The measured value is slightly less than the predicted value of 0.23rad/sec2. At 33" flaps and 80 knots airspeed,ayawingaccelerationof0.22rad/sec2 was achieved at 66% rudder deflection and is about 75% of predicted.

Maximumyawratesachieved duringtherudder reversal maneuversconductedbefore the 5 1 . At 67" flaps, 69 knots, and SAS off, a yaw rate of empennage modification are shown in figure 0.17 radlsec was measured when using 10" rudder deflection. With SAS on, the maximum yaw rate measured was about 0.18 rad/sec with flaps at 33", 80 knots airspeed, and using 60% of maximum rudder pedal deflection. Higher values of yaw rates are shown inthe figure; however, for thosecases the control deflection was decreased while the yaw rate was still increasing.

Lateral-directionalstaticstability- Thestaticstabilitycharacteristics were assessed byper- formingsteady sideslipmaneuvers. Testresultsarepresented fortakeoff,approach,and cruise configurations in figures 52 through 56.

Data are shown for tests conducted both before and after the elevatodrudder modification.

The aircraft exhibits positive stability about both lateral and directional axes for all configurations tested, and the modification produced no noticeable effect.

Figures 52 and 53 present data for the65" flap condition. Figure 52 shows variation of rudder deflection, wheel deflection, bank angle, and elevator position with sideslip; figure 53 shows the same data as a function of rudder angle and indicates a positive dihedral effect. The data fall within a range of C 1 0 as predicted by wind-tunnel tests, correlating better with C l 0 = -0.004 than with Cjp = 0. The variation of rudder and wheel deflection with sideslip angle is reasonably linear out to +15". At 65 knots the sideslip produced by 10" rudder is about 22% less than the predicted value, 90 knots,theaircraft indicatingahigherdirectionalstabilitythanpredicted.Atspeedsbelow exhibited a low-amplitude directional "snaking" characteristic ( p = + l o to 2") indicating that near of sideslip the directionalstiffnessmaybe very low.Attemptstodocumentany zero degrees nonlinearitieswere not successful. About halfavailable rudder and less than halfavailablewheel throw are required to achieve 15" sideslip at 65 to 70 knots.

Figure 54 presents steady sideslip data for the 30" flap configuration at airspeeds of 65, 90, and 97 knots. The rudder deflection with sideslip is linear through the test range of +15" sideslip; however, the slope of control wheel deflectionwithsideslip is reducedwithincreasinganglesof sideslip and indicate a reduction in the apparent dihedral effect at increasing angles of sideslip. The 97-knotdata were collectedafterinstallation of thepoweredelevatorandthe datafor 75 and 90 knots were from flights made previous to the modification. Since the rudder deflection required for a given sideslip angle is identical for both configurations and because the modification reduced the rudder power, it follows that the directional stiffness was reduced as a result of the modifica- tion. This reduction in directionalstiffness was nearly proportionaltothereduction in rudder power.

In the flaps-up configuration, the reduction in apparent dihedral effect with incr,easing sideslip angle is moremarked, as shown in figures 55and56thanforthe 30" flapconfiguration. At 120 knots, about 15" of control wheel is required at 5" sideslip angle while only about double that is requiredforathreefoldincrease in sideslipangle (15").The flaps-upconfigurationtestdata show approximately a linear variation of rudder deflection with sideslip over the range of conditions tested.

In the STOL landing configuration, 15" of sideslip is achieved with bank angles of about 5" , indicating low net sideforces. Ratios of bank angle to sideslip angle in steadystate sideslip,were approximately as follows:

deg I knots

65 6 65 30 6 75 90 .30 5.6 6 120 5.6 6 150 - The low q5/p at landing approach speeds below 65 knots results in very little lateral acceleration with sideslip. The absence of the lateral acceleration cue requires the pilot to use the sideslip indicator in the cockpit much more as speed is reduced to maintain low sideslip angles.The lowdirectional stability aggravates this problem.

Longitudinal trim change with sideslip is small for d l configurations, requiring a maximum of about 1" elevator deflection change over therange of sideslip tested.

Lateral-directional dynamicstability- Lateral-directionaldynamiccharacteristics were eval- uatedatflapsettingsof 65", 30", and 5.6" withthe roll andyaw SAS onandoff (SAS is automatically off above 100 knots). In addition, evaluations were made in the landing approach configuration with only the roll SAS operating (yaw SAS off) and with only the yaw SAS operating (rollSASoff).Dynamiccharacteristicsaresummarized in table 9, Dutch-rollcharacteristicsare shown in figure 57,androlldampingcharacteristics areshown in figures 58 and59.Figure 60 shows a typical Dutch-roll time history at 67 knots airspeed.

The flagged data points shown in figure 57 are from flights conducted after the elevator/rudder modificationandSAS gearingchange (fig. 14(a)).Atthetimeofinstallation of thepowered elevator, the rudder area was reduced 2.5%, directional inertia increased about 5% and rolling inertia increased about 1% at 178,000 N (40,000 lb) gross weight. Theinertiachangehadonlya small effect on the Dutch-roll Characteristics (SAS off), while the SAS gearing change had a large effect.

As noted in the discussion on static stability, a residual directional snaking motion was common at speeds below 90 knots even with SAS on. The SAS gearing was changed to increase the p damping which markedly improved the pilots' opinion of the directional characteristics at landing approach speeds by reducing, but not eliminating, the snaking motion.

At the landing approach flap deflection of 67", the Dutch-roll damping with SAS off is low with a damping ratio of between 0.1 and 0.2. The Dutch-roll period is 6 to 7 sec and is close to the predicted value at60to70knots.The measured Dutch-roll period is slightly less thanthat predicted with flaps 33" at about 100 knots airspeed, indicating a trend toward greater directional stiffness than predicted at higher speed with takeoff flaps. With flaps up, the Dutch-roll period and damping are about as predicted. The oscillation is primarily a yawing motion with I@I/IpI about 1 .O.

With SAS on,the Dutch-roll damping ratio was about 0.3 (flaps 67") before the SAS modification, and about 0.45 (flaps 62") after the modification. The Dutch-roll period with SAS on was about 7secbefore andabout 9 secaftermodification. With either SAS configuration,the roll-to-yaw ratio (I@ I/DI) is about 0.75 with flaps extended (65') and 60 to 70 knots airspeed.

In the landing condition, with the roll SAS off, there is little effect on the Dutch-roll; however, with the yaw SAS off, the Dutch-roll is similar to that when both roll and yaw SAS are off.

With flaps 33", the Dutch-roll damping ratio with SAS off is about 0.1 at 78 knots, but with SASon damping is increased to asatisfactory level ofabout 0.3 at 79knots.The Dutch-roll damping ratio was increased to 0.7 as a result of the SAS modification. With 30" flaps the Dutch-rollperiod is little changed by SAS exceptthatthe period is about2 seclongerforthe modified SAS configuration. At 78 knots the periodis 6.8 sec with SAS off.

In the cruise configuration, Dutch-roll characteristics are satisfactory with period and damping near predicted values. The damping ratio is about 0.15 and period is 4 to 5 sec.

Figures 61 (a) through 61 (g) present typical time histories of Dutch-roll oscillations with SAS on and off, for the various configurations tested.

Time histories showing aircraft spiral characteristics are shown in figures 62(a) and 62(b). The spiral time to half or double amplitude is given in table 9 for the various configurations tested.

With SAS off at 67" flaps, and nozzles up, the aircraft is spirally unstable as predicted, with time to double amplitude of about 6 sec at 60 to 75 knotsairspeed. With nozzle angle increased to 90",time to double amplitude appears to increase;measured times varied from about 8 t o more than 20 sec. Precise spiralbehavior was not alwaysdiscernible due to smalllateraltrimoffsets produced by poorcontrolcenteringandatmosphericperturbations. With onlytheyawSAS operating, the spiral mode was similar to the SAS-off condition with a tendency to diverge more 4 to 6 sec. With only the roll SAS operating, the spiral rapidly, the time to double amplitude being stability is positive. Times to half amplitude of 7 to 10 secwere noted at 69 knots. With all SAS on, spiralstability at 65"flaps is neutral to slightlypositive.Increasingnozzle angle to 90" appears again to give asmallincrease in stability. With 30" flaps the spiral is unstablewithSASoff

(T2 - 8 sec at78knots),andthe stability is neutral t o positive with SAS on.Inthe cruise

configuration, the spiral stability is neutral to slightly positive at 135 and 168 knots airspeed.

The airplane roll damping characteristics are summarized in figures 58 and 59 and in table 9.

The roll damping with 67" flaps, nozzles aft, with SAS off is low with a roll time constant of 1 .O sec as determined from roll reversal maneuvers (fig. 36). The apparent roll time constant with SAS on (67"flaps) is 0.45 sec.Withflaps33" the roll timeconstant is about 1 .O sec (fig. 59) and is approximately the same SAS on or off. With flaps up the roll time constant is 0.8 sec at 135 knots and 0.7 sec at 168 knots; both are near predicted values.

Turn coordination- Turn entries were conducted with flaps at 65", 30", and 5.6' andwith SAS on and off. All tests conducted at less than 65 knots were performed after the elevator/rudder and the SAS modifications were made; therefore, it is not evident what effect these modifications had on turn coordination. Based on the Dutch-roll and steady sideslip characteristics, the SAS-off turn coordination was littleaffected by thesemodifications;however,the SAS-on characteristics were undoubtedlychanged.Theturncoordinationdataobtainedareshown in figure 63 as the observed peak sideslip to bank angle ratio (AO/A@) which was measured at the initial entry to the turn. Turn coordination characteristics where AP/A@ is greater than about 0.3 is considered unsatisfactory(ref.16).WithSAS on, the aircraftexhibitssatisfactorycharacteristics at speeds above 65 knots. Between 60 and 65 knots (SAS on) the turn coordination is marginally acceptable withthe SASgearing used in thetests.There isconsiderablescatter in this region withratios increasing to about 0.4 at 60 to 65 knots. With SAS off, the turn entry degrades markedly due to adverse sideslip giving AP/A@ ratios at speeds less than 80 knots which are above the maximum level of 0.3 generally considered acceptable.

The time history of a turn entry at 68 knots (fig. 64) shows the large adverse sideslip generated on initiation of the maneuver. Yaw rate lags roll rate by about 2 sec. The large sideslip with bank angle andthe lag betweenbank angle andturnrategreatly increase the pilot'sworkloadwhen maneuvering laterally during low speed approaches withoutSAS. With SAS on, turn coordinationis much improved, as shown in figure 64. Yaw rate now follows roll rate with a small lag, and adverse sideslip is reduced.TheSASreduces Ap/A@ ratios toabout 0.3, a value the pilotsconsidered satisfactory. The stability augmentation improves the turn coordination by deflecting the rudder proportional to roll rate. Adverse yaw due to lateral control is negligible in both examples.

At 33" flaps with SAS off, some difficulty was encountered in setting up initial steady-state conditions, with p varying * l o about a steady bias of approximately - 2". The Ap/A4 ratio data in figure 63, therefore, showed scatter with an average value of about 0.37 at 77 knots. No adverse yaw due to lateral control deflection was evident. But yaw rate lagged roll rate by 1 to 2 sec due to adverse sideslip, thus giving unsatisfactory turn coordination overall. With SAS on, adverse sideslip and yaw rate lag were reduced and turn coordination was considered satisfactory.

In the flaps-up configuration (with SAS automatically off above 100 knots), AO/A@ has a value of about 0.1 at 165 knots and 0.13 at 135 knots. In flaps-up turn entries, no adverse sideslip was evident and yaw rate followedroll rate without lag, indicating satisfactory turn coordination.

Engine out lateral-directional control- Time histories of control and airplane motion recorded atthetimeofsimulated enginefailuresarepresented in figures 33(a)and33(b)foralanding approach 165,000 N (37,000 lb) and takeoff at 210,000 N (47,000 lb) gross weight.

In thelanding-approachsimulated enginefailure case shown in figure 33(b),thestarboard engine was abruptlythrottled back to idle from95% rpm whileflying at70knots airspeed.

Approximately 3 seclatertheport engine was advancedfrom 94.3% to 99% rpm. The engine nozzles at the time were deflected 65" to 70". As canbeseen in the time history of wheel and rudder positions, little rolling or yawing is experienced until the port engine thrust is increased. At that time a rolling disturbance to the leftis evident. About 15" of right wheel is used to counteract this roll and the rudder does not exceed 5" until 9 sec after the engine is throttled back. As the engine decelerates to idle, which requires about 25 sec, the airspeed is increased to 80 knots and the nozzles broughtupto 12". Atthiscondition 20" t o 25" ofright wheel and 8" leftrudderare required fortrim.Theangleof sideslipincreased to about 9" about 1 1 sec after the engine was throttled backand the sideslip was thentrimmednearzero as theairplanereached 80 knots airspeed .

The simulated engine failure time history for the takeoff configuration is shown in fig- ure 33(a). The starboard engine was throttled back from 99% rpm to idle about 4 sec after lift off.

The airspeed at the time the throttle was retarded was 90 knots. Four seconds after the starboard throttle was retarded the port engine was advanced to about 101% r p m from its takeoff setting of 99%. Immediately after the throttle was retarded, a yaw to the right was experienced; it required about 6" of rudder for correction, and after the initial disturbance only about 3" of rudder was required for trim. In contrast to the lateral control required for the landing configuration case, the control required for 30" flaps appears to be insignificant, with a small amount of left wheel used to establish a 3" bank angle to the left. Thesideslip is near zero until the left roll attitude is established where it is erratic around about5" left sideslip.

Pilotsconsideredthelateral-directionalcharacteristicswithasimulated enginefailure to be acceptable for an emergency condition.

Lateral-directional control in the stall- The lateral-directional behavior of the airplane in the region of the stall is summarized in table 7 for landing approach, takeoff, andcruise configurations.

The lateral-directional characteristics in the stall are considered mild except with 66" flaps, nozzles up, and one engine at idle. In this case, nearly full lateral controlis required to keep the wings level.

Figures 34(a) through 34(d) show the control used in the stall for various configurations.

75", there is somelateraldirectional In the landingapproachconfigurationwith nozzles at instabilityapparent as the stall is approached,but it is easily controlled using about 5" rudder deflection and a maximum of about 20" wheel deflection. At the stall there is no additional rolling or yawing evident in a typical time history (fig. 34(b)).

In the takeoff flap configuration with engine rpm at 99.676, there is little roll or yaw as the stall is approached, as shown in figure 34(c). Less than 3" rudder and about 3" wheel deflection is used controlling the airplane at the maximum angle of attack. The airplane was not stalled at this condition because of the extremenose-high attitudes encountered.

With flaps up and with 90% rpm, a yaw and roll to the right is encountered at the stall (see fig. 34(d)). About 25" of control wheel deflection and less than 2" of rudder deflection areused for control in the stall. About 9" left sideslip was experienced at angles of attack 1" to 2" above the stall angle of attack.

CONCLUDING REMARKS A flight investigation of 200,000 N (45,000 lb) augmented jet flap STOL airplane has shown that the stability, control, and handling qualities are satisfactory for VFR tasks over the lowspeed operating envelope (60 to 160 knots). Lateral and directional stability augmentation were required t o improvetheDutch-rolldamping,spiralstability,rolldamping,andturncoordination.The longitudinal characteristics were satisfactory for VFR tasks but some form of augmentation would be required for IFR tasks at the low speed required for STOL operations.

Ingeneral,theairplaneexhibitedconventionalflyingqualitieswith the exception of flight- pathcontrol in the landingapproach. With thrustvectoringcapability, an additionalcontrol is available tothe pilotforflight-pathmaneuvering.Thepilotsused thethrustvectorcontrol to establish landing approach flight-path anglewhile the engine throttles and elevator wereused t o establish the desiredairspeedand angle ofattack. When thrustvectormodulation is used to control small deviations from the flight path in the landing approach, the airplane response is similar to that ofconventionalairplanes.Controloflandingapproach flight path by use of theengine throttles produced the tendency for unconventionally large deviations of airspeed from trim.

The stability and control characteristics of the airplane with one engine inoperative (simulated) were acceptable for an emergency condition in landing approach and takeoff. Cross ducting of the augmentor and aileron BLC air provide characteristics that result in satisfactory controllability for the one-engine inoperation condition.

Apoweredelevatorcontrol was required to providesatisfactoryfeelcharacteristicsandfull elevator deflection capability at lowairspeeds. This improved pitch control is particularly noticeable in the stall recovery (landing approach configuration) and in the landing flare.

Ames Research Center National Aeronautics and Space Administration Moffett Field, California 94035, November 1 I , 1977 REFERENCES 1. Wick, Bradford H.; andKuhn,Richard A.: TurbofanSTOL Research at NASA. Astronaut. & Aeronaut-, vol. 9, no. 5, May 1971, pp. 32-50.

2. Whittley, D. C.: The Augmentor Wing: A New Means of Engine Airframe Integration for STOL Aircraft. ICAS Paper 64574, presented at FourthICAS Conference, Paris, Aug. 24-28,1964.

3. Middlebrooks, J. E.; Tinney, H. C.; and Whittley, D. C.: The Evolutionary Development and Current Status of the Augmentor Wing Concept. Paper 700812 presented at S A E National Aeronautics and Space Engineering and Manufacturing Meeting, Los Angeles, October 5-9,1970.

4. Kelley, G. S.; andGerend, R. P.: PropulsionSystems for Commercial STOL Aircraft. AIAA Paper 71-746, A I A A / S A E 7 t h Propulsion Joint Specialist Conference, Salt Lake City, June 14-18,1971.

5 . Quigley, H. C.; Sinclair, S. R. M.; Nark, T. C., Jr.; and OKeefe, J. V.: A Progress Report on the Development of a n Augmentor Wing Jet' STOL ResearchAircraft. SAE Paper 710757, NationalAeronauticandSpace Engineering and Manufacturing Meeting, Los Angeles, September 1971.

6. Ashleman, R. H.; and Skavdahl, H.: TheDevelopment ofan Augmentor Wing JetSTOL Research Aircraft (Modified C-8A), Vol. I: Summary, NASA CR-114503, 1972.

7. Spitzer, R. E.; Rumsey, P. C.; and Quigley, H. C.: Use of the Flight Simulator in the Design of a STOL Research Aircraft. AIAA Paper 72-762, August 1972.

8. Innis, Robert C.; and Anderson, Seth B.: Comparisons of Simulator and Flight Results on Augmentor-Wing Jet STOL Research Aircraft. NASA SP-320, 1972, pp. 283-290.

9. Whittley, D.C.: The Aerodynamics of High Lift Illustrated by Augmentor-Wing Research. CAS1 Paper 72/20, presented at 12th Anglo-American Aeronautical Conference, Calgary, July 1971.

10. Skavdahl, H; andPatterson, D.H.: TheDevelopment ofan Augmentor Wing Jet STOL Research Aircraft (Modified C-8A). Vol. 11: Analysis of Contractor's Flight Test. NASA CR-114504, 1972.

11. Quigley, Hervey C.; Innis, Robert C.; and Grossmith, Seth: A Flight Investigation of the STOL Characteristics of an Augmented Jet Flap STOL Research Aircraft. NASA TM X-62,334, 1974.

12. Glende, W. L. B.: Design of a Powered Elevator Control System - Final Report. NASA CR-114727,1974.

13. Wingrove, Rodney C.: Estimation of Longitudinal Aerodynamic Coefficients and Comparison with Wind-Tunnel Values. NASA TN D-7647,1974.

14. Spitzer, R. E.: Predicted Flight Characteristics ofthe Augmentor Wing Jet STOLResearchAircraft. NASA CR-114463,1972.

15. Wingrove, Rodney C.: Parameter Estimation of Powered-Lift STOL AircraftCharacteristicsIncluding Turbu- lence and Ground Effects. NASA TM X-62,382, 1974.

16. Innis, Robert C.; Holzhauser, Curt A.; and Quigley, Hervey C.: Airworthiness Considerations for STOL Aircraft.

NASA TN D-5594,1970.

17. NATO-AGARD: V/STOL Handling-Qualities Criteria. AGARD Report 577, Dec. 1970.

18. Allison, R. L.; Mack, M; andRumsey, P. C.: Design EvaluationCriteriaforCommercialSTOLTransports.

NASA CR-l14454,1972.

19. Military Specification - Flying Qualities of Piloted Airplanes.MIL-F-8785E!(ASG), August 1969.

20. TentativeAirworthinessStandardsforPoweredLiftTransportCategoryAircraft.Part XX, FederalAviation Administration, August 1970.

21. Military Specification - Flying Qualities of Piloted V/STOL Aircraft. MIL-F-83300, December 1970.

22. Condit, P. M.; Kimbrel, L. G.; and Root, R. G.: Inflight and Ground-Based Simulation of Handling Qualities of Very Large Airplanes in Landing Approach. NASA CR-635,1966.

TABLE 1. RESEARCH AIRPLANE CHARACTERISTICS Weights, N(1b) 21 3,000 (48,000) Maximum gross Maximum gross (STOL takeoff) 200,000 (45,000) Maximum landing (STOL landing) 191,000 (43,000) Maximum fuel 62,000 (1 4,000) Areas Wing area, total including ailerons flaps and 10.3 1 m2 (1 1 1 ft2 ) of fuselage, m2(ftz ) 80.36 (865) 17.38 Wing flap area, projected, including ailerons aft of wing line, m2 (ft2 ) (187.10) Total aileron area aft of hinge line, including trim tab, m2 (ft2 ) 4.30 (46.30) 21.65 Horizontal tail area, total, m2 (ft2 ) (233) Elevator area aftof hinge line, m2 (ft2 ) 7.57 (81.5) 14.12 Vertical tail area, total, m2 (ft2 ) (152) Rudder aft ofhinge line 2.65 Fore, m2 (ft2 ) (28.5) Trailing, m2 (ft2 ) 2.79 (30) Dimensions and general data Wing Span, m(ft) 24.00 (78.75) Root chord, m(ft) 3.83 (1 2.58) 2.36 Tip chord, m(ft) (7.74) 3.78 Reference aerodynamic chord, m(ft) (1 2.4) Aerofoil section NACA 64, A41 7.5 (MOD) Root Tip NACA 632A6 15 (MOD) Sweepback at 40percent chord, deg 0.0 Dihedral, outer wing only, deg 5.0 (Note: Wing taper and dihedral each start 5.36 m from plane of symmetry.)

Aspect ratio 7.2 Ailerons Span, m(ft) 3.5 1 (1 1 S O ) Chord aft of hinge line, m(ft) .61 (2.01) Distance from plane of symmetry to centroid of aileron, m(ft) 10.27 (33.70) 20.0 Aerodynamic balance, percent Spoilers Span, m(ft) 3.44 (1 1.30) Chord, m(ft) .36 (1.18) Position of hinge line, percent wing chord (average) 62.4 Flaps Span (each side),m(ft) 7.01 (23.0) Chord aft of hinge line, m(ft) .98 (3.2) TABLE1.RESEARCHAIRPLANECHARACTERISTICS - Continued .'; Horizontal tail 9.75 (32.0) Span, m(ft) 2.54 (8.33) Root chord, m(ft) 1.91 Mean aerodynamic chord, m(ft) (6.25) Aerofoil section NACA 63A2 14 (MOD) Root (inverted) NACA 63-21 2(MOD) Tip (inverted) 4.8 Sweep of leading edge, deg 0.0 Dihedral, deg 4.4 Aspect ratio Vertical tail 4.14 (13.60) Span, m(ft) 4.27 Root chord, m(ft) (1 4.00) 2.54 (8.33) Tip chord, m(ft) .+.

-, , 3.48 Mean aerodynamic chord, m(ft) (1 1.41) NACA 63(2 1 5 101 4 (MOD) Airfoil section, m(ft) 22.6 Sweep of leading edge,deg 1.2 Aspect ratio 8.75 (28.7) Overall height, m(ft) 28.44 (93.32) Overall length (with nose boom of 4.88 m) m(ft) 14.1 1 (46.3) Distance, wing MAC, 1 /4C, to horizontal tail MAC, 1 /4C, m(ft) 13.23 (43.4) Distance, wing MAC, 1 /4C, to vertical tail MAC, 1 /4C, m(ft) +2.5 Wing incidence angle, deg

+I .o

Horizontal tail incidence angle (ground adjustable), deg Control surface deflections and rates Flaps 5.6" down to 72" down 4"lsec extension and retraction Conical nozzles 6" t o 104" (down from aft of aircraft) 90" /sec +19" about +35" max droop angle Ailerons (65" flap deflection) 50" /sec 48" Spoilers 1 20" lsec 55% choke gap area closure at 75" flap deflection Augmentor choke 35"/sec Rudder +25" forward segment +25" trailing segment - 5O"/sec Elevator (at 65 knots) - 24" +15" TABLE 1 . RESEARCH AIRPLANE CHARACTERISTICS - Concluded Maximum design speeds, knots Dive speed ( V o ) 180 Maximum operating speed ( Vmo) 160 30" flaps-down speed ( VDO) 120 >50° flaps-down speed ( V ~ 5 0 0 ) 90 Moments of inertia, at 178,000 N (40,000 lb) gross weight With powered elevator, kg-m2 (slug-ft2 ) I , = 361,000 (266,300) Iv = 3 16,000 (233,100) I, = 620,000 (457,300) With spring-tab elevator, kg-m2 (slug-ft2) I , = 357,000 (263,300) Zv = 278,000 (205,000) I, = 587,000 (432,900) TABLE 2. ELEVATOR RESPONSE AND FLIGHT-PATH CONTROLFORTHE STOL APPROACH CONDITIONS (65 knots) Subject

Criterion 1

airplane t Value

Reference ~~ ~ ~~ ~ " ~ Elevator control power (FcoL = 180 N = 40 Ib), rad/sec2 0.47, -0.37 X.05 to 0.20 17 Pitch controlsensitivity

" _

(q/6coL), rad/sec2 cm .03 rad/sec2 in.

.08 >.08

Stick force per g, N/g " _

W g <20 to 40 > 3 17

Effective vertical speed crossover time (ti), sec .9 e . 8 Load factor response time constant (7 ), sec "1.5 <1.5 16 %

Load factor per unit control deflection (nz/ScoL), g/cm " -

.04 g/in. .10 -. 1 Load factor per unit angle of attack @,/aF), glrad 1.8 >2.3 18 Flare control load factor,g >1.20 31.20 18 Pitch angle after 1 sec, deg (FcoL = 180 N x 40 Ib) 3.1 17 >2 to 4 ~~ ~~ . ~~ . . .. ~ TABLE 3. ELEVATOR SYSTEM CHARACTERISTICS ATSTOL APPROACH SPEED (65 knots) - - " Subject Crite ric In Value Referencc . _ _ _ _ _ _ _ airplane

I ~-

Elevator deflection from trim at 180 N (40 lb) Fcok del -24, +14 Control column deflection, cm +19, -13 in.

+7.5, -5 17 Elevator to column gearing, deg/cm 1.9 deg/in.

4.8

Force gradient (AS, < lo"), N/cm

10.1 Ib/in.

5.8 Column force at T.E.U. maximum S , , N Ib 42 20 .- " TABLE 4. LONGITUDINALSTABILITYCHARACTERISTICSAT STOL APPROACH SPEED (65 knots, 178,000 N Gross Weight) Criterion Subject

aiiplane Value I Reference

0.1 x 20 Stick-fixed stability (AsJAV,), deglknot G.06 Change of flight path withvelocity (A~/AVE) deg/knot 0 19 Short-period natural frequency ( a , ) , rad/sec 1.03 >.47 21

Short-period damping ratio (c) .9 >1.06 21

" _

.2 Phugoid natural frequency ( a , ) , rad/sec

Phugoid damping ratio (c) .1 x 21

-.8 Change of 0 with velocity (AO/AVE), deg/knot <O 18 TABLE 5. AIRPLANE SHORT PERIOD AND PHUGOIDCHARACTERISTICS Phugoid Shortperiod Weight P, "n 7 P, "n sec sec { rad/sec { rad/sec 1.00 37.1 0.09 0.17 9.1 0.72 184,000 41,400 33.2 .09 .19 16.5 .90 .87 172,000 38,700 33.2 .10 .19 14.1 .90 1.03 167,000 37,500 34.9 0 . I 8 10.1 .86 1.20

178,000 40,000 1

TABLE 6. STABILITY AND CONTROL DERIVATIVES FOR THE LANDING APPROACH CONFIGURATION 6 f = 67" ffF = -2" to 10" v = 14" VE = 66 to 79 knots NH = 94.5% i ' , , = 281,000 kg-m2 (207,000 slug-ft2) CJ = 0.3to 0.4 Stability Flight Predicted nd control measured lerivatives 5.07 4.79 - .68 -.55 - 17.0 -loa - 3.00 -3.45 1.32 1.33 2.22 3.56

" _

.18 1.57 1 .oa - .55 -.52b - 2.05 - 2.0Sb - 29.7 - 350 . .

a Predicted tail contribution only, q + C $ bAdjusted for c.g. location and Cm aF TABLE 7. STALLING CHARACTERISTICS ” Nozzle Engine rpm Stall angle of Stall airspeed Flap deflection Behavior at stall deflection u, deg Gross weight, W attack, aF NH, %

%

knots Buffet (typical) lip deg rightlleft rightlleft N (1b) deg 4- 47 Slight G break, some directional 65.0 76/72 96.6196.9 191,000 (42,900) 27 snaking, pitch-up tendency 64.5 95.0/95.8 193,000 (43,400) 24 51 Moderate Slight G break, yaw and 711 2 roll to right, full fwd.

control used 65.7 Idlel99.6 197,000 (44,300) 18 65 Moderate G break, roll and yaw to 73/72 left, 1at.-dir. oscillation 65.8 9/13 Idlel97.0 187,000 (42,000) 21 63 Mild G break, yaw left, pronounced 1at.-dir. oscillation, full lat.

control required 31 .O 12111 93.0194.0 192,000 (43,200) 23 58 Mild Slight G break, roll and yaw to left 30.3 97.O/Idle 192,000 (43,200) 66 Slight Slight G break, yaw left, roll 718 22 right, 1at.-dir. oscillation 5.8 90.1/90.6 197,000 (44,300) 28 69 Heavy Slight G break, roll and yaw to right 5.6 l o l l 0 97.3/Idle 178,000 (40,000) 26 67 Heavy G break, roll and yaw to right w w TABLE 8. LATERAL-DIRECTIONAL STABILITY AND CONTROL DERIVATIVES

l & f = 65"

W = 182,000 N (40,900 lb) NH = 96.5% CIF = 5.6" u = 87' CJ = 0.47 C = -0.169 'P = -0.502 = 0.436 'lr = 0.272 Cn = -0.937 c"P = -0.727 C p = -1.32 'nr YP C = -0.044 c = 1.12 Y r YP i TABLE 9. LATERAL-DIRECTIONAL DYNAMIC CHARACTERISTICS 'r Roll mode

Dutch roll mode I Spiral mode

-r

Weight, V ' knots SAS C J T ~ , sec N X 1 0 3 P, sec +, rad/sec

t TI / 2 , sec

6f I v I

-.-

.o 165 to 178 0.23-0.35 6 -0.1 5 1.06 1 OFF

" _

169 .34 7 .13 .9 1 Roll on/yaw off

" _ " _

" _ " _ " -

169 .34 Roll off/yaw on 173 6.8 .33 .98 Neut + pos 0.45 .35 ON

_ " " _ " _ " _

.40-.45 173 Neut +pas " - .43-.49 173 to 196 8.8 .46 .80 Neut + pos 74"-85" 62-66

" _ " -

173 to 182 .43 7.0 .14 .90 90" 62

OFF^

" _

-1 .o 178 .15 4.9 .16 1.3

OFF 33" I -10" I 102

" _ " _

1.45 105 178 .21 4.3 .12

" _ " _

78 198 .27 6.8 .10 .94 Neut + pos -1 .o 79 195 .29 6.4 .3 1 1.03 ON

I

" _ " _

.7 1 1.06 30" 77 200 .32 8.4 ONa

I

.16 1.51 Neut + pos .70 168 200 .06 4.2 Not applicable 6" -10" .075 .15 1.27 Neut + pos .80 135 178 to 196 5 .O Not applicable 6" -1 0" a These three conditions from flights conducted after the elevator/rudder and SAS modifications; all others from flights conducted before the modifications.

TABLE 10. LATERAL AND DIRECTIONAL CONTROL CHARACTERISTICS Criterion Subject airplane Value Reference Lateral k7 5 260 16 Control wheel travel atmax deg

" _

Control travel, cm k28.5

" _

in. +I 1.2

" _

Breakout force, N k22.2 k5.0 1/2 to 4 21 lb " -

Force gradient ( 6 w < 20°), N/cm 1.40

1 to 3 17 lb/in. .8 " _

Maximum force at max b, N 20.0

lb 4.5 <20 16 <0.3 for 63% p max Control response (full control), sec .3 17 Directional

" _

Rudder pedal travel, cm k9.9 in. k3.9 2.5 to 4.5 17 Force gradient (k25% rudderpedal

" _

travel), N /cm 45 10 to 35 17 lb/in.

Maximum force, N 43 6 " _ lb <130 Control response (full control), sec 0.7 <0.3 for 63% i max IAHLC I 1. LA I CKAL A l Y U UlKCL 1 IUNAL CUN I KUL PUWCK L'HAKAL'I EKISIICS FOR LANDING APPROACH AND TAKEOFF CONDITIONS Subject airplan Criterion Landing Takeoff Value Reference SAS off SAS on SAS on - - ~ Lateral

6 max, rad/sec2

0.67 0.70 0.60 X . 4 16

" _

.57 9, IS,, deg/cm* .39 .49 deg/inic

1 .o 1.46 1.25 X . 8 22

Time to 30", sec* 2.4 1.75 1.9 e2.4 16

P / S , , (6, < 40")

" _

rad/sec2/cm .042 .043 .037 rad/sec2/in. .lo6 .110 .095 X . 0 7 22 Directional 0.37 rmax, rad/sec2 0.37 0.40 x . 1 5 17 20 20 20 3 2 5 18 @,ax? deg

" _

23 10 to 15 18 &crab (A$* . o h deg

" _

Decrab time (A$ = 1 So), sec 1.6 1.5 <2.0 17

Sensitivityr(Sp < 2.54 cm)

" _

rad/sec2/cm .047 .047 .05 1 rad/sec* /in .12 .13 .12 X . 0 5 17 *Based on a ramp wheel input (ramp time = 0.5 sec) TABLE 12. LATERAL-DIRECTIONAL DYNAMIC CHARACTERISTICS FOR LANDING APPROACH AND TAKEOFF CONDITIONS Subject airplane Criterion Landing Takeoff Value Reference SAS SAS SAS on off " 1 - on .

Dutch roll damping (cud), 1 /sec 0.37 0.15 0.75 X . 0 8 7 16 Dutch-roll frequency (ud), 1 /sec .80 .90 1.06 >.523 16 Roll time constant ( T ~ ) , sec .45 1.1

1 .o <1.4 19

> - .035 16 Spiral time constant (TJ sec "0 -.OS -0 Turn coordination (APIA@) .25 .15 <0.3 16 .65 Dihedral effect (dS,/dO) 1.2 1.2 1.4 X 17

" _

Dihedral effect ( F , 1 7 N 35 35 35 omax Ib 8 8 8 <10 17

I

(a) Landing configuration.

(b) Takeoff configuration.

Figure 1 .- Augmentor wing jet STOL research airplane.

P Spoiler (lateral control) Maximum gross weight = 214000 N (48000 Ib) vM0 (flaps up) = 160 knots Aspect rat io 7.20 deflect to 7 3 O Wingarea 80.4 m2 ( 8 6 5 f t 2 ) 23.9 m ( 7 8 . 7 5 f t ) Span Flap semispan 7.Om (23 ft) ( 7 8 . 7 5 f t ) Leading-edge slats Modified

Pegasus (30.5 f t ) Spey Rolls-Royce MK 801 - SF L k z (27.83 f t ) 23.5m - 4

nozzles (77.33ft) Figure 2.- Three-view drawing of modified C-8A.

Bypass air

Vectorable swiveling

' Bypass air

Main

engine mount mount assembly aft duct

Figure 3.- Nacelle installation.

I, , 1 1 1 I Engine throttles (2) Engine nozzle vector controls (2) Elevator and lateral trim switch (a) Overhead console.

Sideslip -

Angle of ati Airspeed (b) Pilot’s instrument panel.

Figure 4.- Cockpit interior.

Outer Upper duct nozzle Wing chord plane duct Lower nozzle Choke Figure 5 .- Air distribution system.

= zero

-40

0 5 0 - 5 - IO

Elevator trim, deg

-

Li

-

””” ” 8 4 , LL

-TFriction

Trim 8 e =Oo

CI Q)

L 0 ’ - z 0

*

/Friction

c

“-

’ - “

- “

- 20

5 -4

-

‘Break-out force

-

-8

I I

-40

0 40 80 120 160

Airspeed, VE, knots

Figure 6.- Elevator control breakout and friction forces.

- m

Q) W \

I t

I O

i-

-

'20 40 60 80 1 0 0 120 140 160

Airspeed, VE, knots

Figure 7.- Elevator control feel force gradient.

I I I I Ill1 I I I 1 I I l l 1ll11mlII IIll1111l1 II

Column length= 82.7 crn

Note: 180 knot curve offset by -I" s e for clarity

-24

-20

-I 6

u l -12

Q) U CI Q)

* CI -8

c

.-

t

- Q) -4

u- Q) U

t z o

-

W

1 2

1 6

-

8 - 4 0 4 8 1 2 1 6

Coiumn deflection, BCoL, deg

Figure 8.- Elevator to control column gearing.

1 . .

-

-30

I

Control column position limit

I Maximum

Q) 0 - 2 0

-a

allowable

a

/Trim limit

design tail load

4 8

I

” 0 -10

t

Maximum

Q)

-

operating +

P

Q)

. ~~ - speed

L 0

t

>

Q)

-

” IO

Blowdown

0 40 80 I20 I 6 0

Airspeed, VE, kn o ts

Figure 9.- Elevator control and trim authority.

/-

20 30 F

/ - -

l.F

- I 0

L Q)

s -20

i

E bL'

-3 0 u

-10 -5 0 5 1 0

Rudder pedal position, 8 p , cm

-250

-500

-25 -20 -15 - 1 0 -5 0 5 IO 1 5 20 25

Rudder deflection, 8 , , deg

Figure 10.- Directional control characteristics, pedal force and gearing.

forlateralcontrol

~ C H =65% closed

for l i f t dump

Lateral control f 19' about the droop

65'

position with flap deflection of

Note: Outboard choke - lateral control and lift dump

Inboard choke- lift dump only

Figure 1 1 .- Augmentor flap and lateral control geometry.

m

a

-

m

2 30

P

-g 20

c

L

a

-

I O

”

a

I

0 IO 20 30 40 50 60 70 80

Flap deflection, Bf, deg

(a) Aileron droop program.

Figure 12.- Aileron mechanical characteristics.

1 6

-8

-16

-24

-32

-40

-100 -80 -60 -40 -20 0 20 40 60 80 100

Control wheel deflection, Bw, deg

(b) Control wheel to aileron gearing, SAS on.

Figure 12.- Continued.

5 1

.- 50

a,

5 20

E a

.-

x

IO

0 20 40 60

Flap deflection, 8f, deg

(c) Maximum choke deflection program.

Figure 1 2.- Continued.

I O

I

-80 -60 -40 -20 0 20 40 60 80

Control wheel deflection, BW, deg

(d) Choke and spoiler gearing, SAS on.

Figure 12.- Concluded.

60 80 100

Control wheel angle, &, deg

Figure 13.- Control wheel forces.

Bank a Directional SAS servo Gain= 0.4 (first 78 flight test hours) Gain = 0.8 (remaining flight tests) Authority limited to 8, = * 5" SAS servo gearing 10" rudder deflection per volt (a) Directional axis.

Figure 14.- Stability augmentation (normal mode).

5 5 Flaps > 4 0 " ,

I .50 1

Roll rate, deglsec b

0 +

Flaps ~ 4 0 "

+ Roll dumping

augmentation

b 0.75

I Lateral series

SAS servo

I . 7 5

--%"

n

Spiral stability augmentation

Electricu I

* knots

gearing

Authority limited to an equivalent wheel deflection (&) of +, 20'

SAS servo gearing 1 0 " (equivalentwheeldeflection) per volt

(b) Lateral axis.

Figure 14.- Continued.

Pi lot's co ilot's

wheel w I: eel

I Mechanical

- gearing

I"

I

Central TO

I - - 4 control 1

+ lateral

I"

actuator controls

I

I

I

I

I

Hydraulic

VE '100 knots &

I

pressure

I

I

'f -

Limiter Electrica I

gearing " Quickener

Lateral S A S

command (Equivalent (c) Lateral control quickener.

Figure 14.- Concluded.

!

"" Center of gravity variation

""

Aft structural limit

"

Aft aero limit

"-

Forward aero limit

Figure 15.- Variation in center-of-gravity location with gross weight. Gross weight changes with fuel only

I O

-10

-10

0 5 1 0 1 5

t, sec

Figure 16.- Response to elevator reversal, 66 knots V,r, 64.8" 6 ~ , 71" v , W = 182,000 N, CJ = 0.39.

knots deg 65 55-58 65 70-74 30 74-78

J

.8 x lo3

"" . 6

i

t

(a) Variation of normalized angular acceleration with incremental elevator deflection.

Figure 17.- Longitudinal control effectiveness with powered elevator.

- Airspeed, VE, knots - 1

*4 I + "

.2

f

0.2

0.4

"t-

-. 6

1 2 8 4

1 6

(b) Longitudinal control power forvarious airspeeds.

Figure 17.- Concluded.

. .6

- Spring tab elevator

.4

r ad/sec

Powered elevator

N/m2

.2

0 - 4 -8 - I 2 - I 6 -20

Elevator increment, As,, deg (measured from trim)

Figure 18.- Longitudinal control effectiveness.

NH, *

v, Weight, CG, Altitude, af,

Y O O10 e m N

d e 9 de9

66.5 90.2 190.7 I 0 30.9 2 IO0 203000

30.5

65.0 8 92.8193.3 2300 I93000

30.6 2200

64.6 IO 95.0195.8 195000

66. I I 2 99.5199.8 30.4 1900 186000

-6

-4

-2

-5 0 5 IO 1 5 20 25 30

40 50 60 70 80 90 100

Angle of attack, QF, deg

Airspeed, VE, knots (a) Nozzle deflection 10".

Figure 19.- Variation of longitudinal static stability withengine rpm and nozzle deflection, flap deflection 65" CG, Weight, Altitude,

O/O E m

N

h

3 0 . 4 2000 189000 0

30.5 2200 190000

A

3 0 . 6 1700 194000

0 197000

1 8 0 0

, 3 0 . 7 ,

A

CT

201000 3 0 . 8 1 8 0 0

Q)

-15

c

2 -10

c "

.- 0 -5

t

P o

Q) B t

> IO

2 40 60 80 1 0 0 - 1 0 0 IO 20 30

W

Airspeed, VE, knots Angle of attack, a ~ , deg

(b) Nozzle deflection 73".

Figure 19.- Continued.

Weight, Altitude, I/, c G,

Bf Y

NH 3

m % E

N

%

deg deg

0 64.6 I O 95.0/95.8 I95000 2200 30.6

27 94.9/95.5 I96000 2300 30.7

65. I

65 .O 5 5 95.6195.9 I98000 2300 30.7

65. I 95.7196. I 1 9 I 000 2300 30.5

2 100

64.5 8 8 94.8/95.5 I87000 30.3

67.0 102 94.2/95. I I86000 30.0

-8

co

=* -6

'F -4

u Q)

w- - -2

Q)

u o

W

1 I I 1 I I 1

40 50 60 70 80 0 5 IO 1 5 20 25 30 35

Angle of attack, aF, deg

Airspeed, VE, knots ( c ) Engine rpm approximately 95%.

a !

u l Figure 19.- Continued.

"

I 0 j 7 2 j ldleI99.6 j 197000 j 2100 i 30.7 I

A

73 ldle/96.5 186000

I O 2300 30.4

Idle/97 186000

1 65 1 ldle/96 1 2300 1 1

1 177000

-I 0

Q)

'0 -8

c .

Q)

m -6

n

c

.2 -4

t Q)

u- -2

Q) I

8 0

c

Q) ' 2

w

40 60 80 I O 0 I20 0 IO 20 30

Angle of attack, aF, deg

Airspeed, VE, knots

(d) One engine simulated failed.

Figure 19.- Concluded.

Altitude, Weight, v, NH, CG, m % E % deg IO 90. I /90.5 I96000 2300 30.6 3 1.2 12 9 1.6/92.0 I86000 2400 30.3 30.8 1 0 93.2193.8 191000 2400 30.5 I O 94.4/ 94.9 I89000 I 900 30.5 30.5 1 0 96.4/96.9 I96000 2200 30.9 28.6 1 0 99.5/99.7 202000 2100 30.9 " TO -25" 8 e at 28" U F

- 12

-10 -8 -6 -4 -2 40 60 80 100 120 140 0 5 IO 1 5 20 25 30 Airspeed, VE, knots Angle o f attack, U F , deg (a) Nozzle deflection 10".

Figure 20.- Variation of longitudinal stability with engine rprn, flap deflection 30'.

I Wetht, I Altitude,

m

0 192000 I400 3 0 . 5

I 0 1 180000 1 1700 1 3 0 . 1 I

I I I 1

40 60 80 100 120 0 5 IO 1 5 20 25

Airspeed, VE, knots Angle of attack, a ~ , deg

(b) Nozzle deflection 1 O", one engine simulated failed.

20.- Concluded.

Figure

NH, I Weight, I Altitude, I CG, I

o

30.9 1400 265000 90.3

2100 I97000

30.4 94.3

A

2100 201 000

30.8 97.4

I 0 I 90;4 I I98000 I 2700 130.7 1

- -20

6 8

-I 5

-10

-5

180 0 IO 20 30 40

20 60 100 140

Angle of attack, C~F, deg

Airspeed, VE, knots

(a) Nozzle deflection 10'.

Figure 21 .- Variation of longitudinal stability with engine rpm, flap deflection 5.6".

Weight, I Altitude, 1 CG,

N O/O E m

30.0 1600 178000

CT Q) U

” -20

-15

- 1 0

-5

60 100

I40 I80 0 IO 20 30

Airspeed, VE, Angle of attack, a ~ , deg

w knots

(b) Nozzle deflection lo”, one engine simulated failed.

Figure 21 .- Concluded.

............................... ~~.~ .................

. . . . . . . . . . . / a . . . . . . . . . . . . . . . . . . . . .

&,>. 1.

. . . . . . . . . . . .?.

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . .

VE, knots

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

" 1 . 1 . . . . . . . . . . . . , . . ' . , Y .... .> ....... . . . . . . . . . . ' ' . I .

. * . . ' : ; . ; ; : A : . .-.. , " I > : : . . . ; , , ... . . . . . ... . . . . . . . . ...................... . '..'

".:."

. . .

,;,v .... -

1' ;.': '.., " . . ,

I .o I

. . . p,; . ..,. ::': ; . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . -.,+,;.. . . . . . . . . . . . . . .

.9 . . . . . . . . . .: . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.................................................................... :._.:..;: 20 seconds ...:..:...,.; . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

c w . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-. .. ..... I . .

. . . . . . . . . . . . . . . . . . . . . .

~::.~~:;~'-;~~,~" ~...~~~,..~.;~~~,"~:.~".~,.;.. ::;<;,; . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

- .........................................................

-. 25

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

, . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . ." ....... - ......... - ...... - . . .

c

,-- "- "_

0 -__ . . . . . . . .

I-.

, . . . . . . . . " ...........

"-;.-+="-.~ . . . .

8, deg . . . . . . . . . . . . . . . . . . . . . . . . . . .

...... . . . . . . . . . . . . . . . . . . . . . . . . . . .

-10 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . , . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. .

Figure 22.- Elevator pulse - flaps deflection 67', nozzles deflection 77", 65 knots.

"-

Constant QF contours """

Constant 8 contours

-*-.- C L , , , ~ ~ contour

Flopplacardspeed

-5

I

i

.- 0-20

G

i

-30 - 2 5 T 30 40 50 60 I/: 70 80 : 90

Airspeed, VE, knots (a) Nozzle deflection constant at 90".

Figure 23.- Landing operational envelope, flap deflection 65', W = 178,000 N, sea level

standard day.

"-

Constant a F contours

"""

Constant 8 contours

-.-.-

contour

C~~~~

Flap placard speed

I 0 4

Airspeed, VE, knots

(b) Engine rpm constant at NH = 94%.

Figure 23.- Concluded.

. . .

..........

"""_. . . .

"""_ ~ - ......

. . . . . . ".~. - . "~ ..

. . . . . . . . . .

. . . . . . . . . . . .

. . . . . . .

. . . . . . . .

. . . .

. . . . .

. .

. . . .

! : . . " . .

. . . . . . . . .

c " . . . . . . .&""".T . , . "?"

"-~"" :.

."

-.,"-., . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . 20 seconds : -

. . . . . . . . . . .

too . . . . . . . .

. . . . . . . . .

. . . . . . . . . . .

. . . . . . .

. . .

~- .. ".

VE knots

. . . . .

. . . . . . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . . . . . . .,- . . . .

. . . . . . . . . . . .

: : : : . : : . : : . . . . . . . .

. . . . . . . : : : , . . : I . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . - .

O r . . . . . . . . . . . . . . . . . . . . . . .

L : ; ; . . .

. . . . . . . . . . . . . . ~ . . . . . . ."

. . , . . . . . . . :.., . . . . . .

. . . . . . . . . . . . . . . . . . .

I . . . . . . . . . :,....

I .5 . .

7- . .

. . . . . . . . .

I .o

. . . .

. . . . . . .

I50 . . . . . . . . .

~.

. . . . . . .

. . . . . . . . . . . . . .

too

. . . . . .

...... . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Altitude, m - . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Figure 24.- Landing approach time history, 6 f = 64.7", W = 193,000 N.

C G, 8fv Weight, O/O 'c N de9 30.5

66 I 9 I O 0 0

30 30.3 I86000

32 30.6 I 920 00

6 30.4 I89000

-

67 Var.

I74000 29.5

67 I95000 30.3

33 I84000 29.9 29.8 33 I82000 -30 Elevator deflection Per 9 ,

- 20

Stick force perg,200

r f

AFcoL 100

- 9

I I

8 - Load factorper unit angle of attack, n 4 - , L i r e d i c + ! d (W=189000N) g/rad n U 40 60 80 100 120 140 160 VE, knots Figure 25.- Maneuvering stability summary.

s f = 29.5O

Weight = I 86000 N

C G =30.3%E

u = g o

NH = 92 . W o

Q) U LL U Z

i

.

LL

Normalload factor, AZ/g

Figure 26.- Wind-up turn, flaps 30°, 90 knots, powered elevator installed.

" " _ Computed for W=Ii'4OOON, C ~ z 0 . 4

1 :

.3

.2

. I

4 0

z l

- . I

- .2

-. 3

(a) Airspeed 63 knots.

Figure 27.- Variation of vertical acceleration with changes in angle of attack, 6 f = 65", v = 60" t o 90".

h N

c

a

I

' Weight, N

-.2

/

0 965000

/

#

- 0 3

/ - - E l

176000

-a 0

"IO

-5

5 IO

(b) Airspeed 68 knots.

Figure 27.- Concluded.

.........

0 . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. .

_ . . , . a .- 4 -10 . . I + , A . _ _ . . ( : : : 1 0 seconds : : : ; "" . .

. . . . . 1 . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . " , . , , .

.... ........ I . , . . .

I 70 4 . . . . . . . . . . . . . . . . . . . . . . . . ..

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . K-.?%.y . . . . . . . . . . . . . . . .

* . -

. . . . . . . . . . . . . . . . . . . . . . .

. i . . . . . . . . . . . . >". . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . b" c

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

. .

. . . . . . . . . . . . . . . . . . . . . . .

" . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. ( . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

I .2 . . . . . , , , . &":%\-, . . . . . . . . . . . . . . .

. . . . . . . . . . . . . + : . . . . ~. . . . . . . .

. . " . < . , . > ." -.- - "" & " " . . * "& , . " . . , - ; . * ' ... .- -: , . . .

.....

*-2 .",.. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

/++:: . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

.8 . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

.....................

...... " ........ - ................ ....

........................ . - . . . . . . . . . . . . . . . . . . . . . . .

IO0

_ _

........................

- .

. _ 0 K

FCOL ' N

..

..

. _ ...

.2 . .

. .

& . .

. .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . .

. . . . . . . . . . . . .

. . . . . . . . . . . Q F : : ' : : : : : . .

. . . . . . . . . . . . . . . . " ' Q F

-. 2 .......

- .

. . . . .

/".

' /"'"" _-. - . . . . . . , . > . . . . . . 4.. . z c . , . , . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

"#.

IO - , . - / - . " J

.-

. . . . . . . . . . . . "-F- A- 5 ' : . . . . . . .

. . . . . . . . . . . . / . . . . .

- . _r . " 0 . .

. . . . - : . . . . . e r . . . .

. . : : . . . : : : : . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . ji: . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

-10 .......

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

IO . . . . . . . . . . . "--,.. . .

. . . . . . . . . . . . . . . .

.-. ,x- . _- k

.. -~ " 1 : , .

-./"--",-- . .

" > i . . . . . . . . . . . . . . . .

, , . , , , , , -:, . : . .

. . . . . . . . . . . . . . . . . . . . . . .

+ : . : : . . . . . . . . . . . . . . . . . . . . . .

-10 . .

. . . . . . . . . . . . . . . . . . . .

......

IO . .

. . . . . . . . . . . . . . . . . . . . . .

. .

. . . . . . . . . . . . . . . . . . . . . .

. .

. . , . . . . . .p: . . . . . . . . .

, .

. .

. . . . . . . . . . . . . . . . . . .

" " '.,~""iL- r= 0 . .

. . . . . . . . . . . . . . . . . . . . .

. .

. . . . . . . . . . . . . . . . . . . . . .

. .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

....

. . . . .

-10

I/ = 88"

u = 88"

W = I86000 N

W = I74000 N

(a) Airplane nose-up steps.

Figure 28.- Pitch attitude steps.

0 ... ............ ~ ......

. . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

: : I : . . . . . .

L_, ' " - -/-

- 1 0 . . . . . . . . . . . . . . . . . . . . . .

. . . . .

r. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . _ . _ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . :/.+ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

......................

. . . . ~ . . . : : . . .

/.- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -/.. ....

. . . . . . . . . . . . . . . .->#A -! . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Y%"K-.e . . . . . . . L - : : : : : : : / - ; I : : : : : : . . . . . . . . . . . . . - i . . . . .

I - . . . . . . . . . . . . . . . . . . . . . . . . _ . . _ . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

I . . . . .

. . . . . : : 1 0 seconds : : , . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . I : 1 . 2 b.;-;.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. .

. . . . .

. .

. . . . . . . . . . . . . . . . . . . . . . .

. - ZN.

. . . . . . . . . . . . . . . . .,< " ! ? - - . ; ? - ' ? - . . . . . . . 8 . . . . . . . . . . . . . . . . . . . . . .

. . , . . I : : . : . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . : . . . . . . . . . . . . . . . . . . . . . . .

_" .. +.-. .- " .

-

. . . . . . . _ . . . . " _ _

".", IO0 ..........".... ~ ......

. . . . .

_ ...................... . . . . .

T ; . . . .

. . . . .

FCOL, N

. . . . . ................_.... ..

. . . . . . .......................

.......................

1 . : : 1

_ . . . , . . . . . . . . . . . . . . . . . . . ;-iT . . . . _ . . . . . . . . . . . . . . . . . . . . . .

.2 . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . .

k : : : : : : : : : : : : : : : : . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . &

-

"7- __

-

q, radlsec o

. . . . . _ . . . . . . Y. . . . T L . , ' . . . . . . . . . . .

. . . . . . . . . . . . I . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . _ . . . . . . . . . . . . . . . . . . . . . . F . . . . . . .

-. 2

..... I. ._ .................................................. .I .....................................

. _ _ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

IO

. . . . . . . . . . OF

. . . .

. . . . . . . . . . . . . . . . . . . . . . I

-

- : . . . . . . . . . . . . . : ~ : - 6 : -

=--- . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . _ . . . . . . . . . . . . . . . . . . . . . .

-10 ............. , ..................................

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . :i . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . _ . . . . . . . . . . . . . . . . . . . . . .

1 0 . . . . . . . . . . . . . . j h ; . . . . . . . . . . . . .

" . . . . . . . ! . - ? I " ? . . . . . . .

. . . . . . . . . . . . . . . . . . _ . . . . . . . . . . . . . . . . . . . . . . . _ _ . . _ _ -10 I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

" ............................................... ....... .....I IO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . r ; . . . . . . . . . . . . .

? " . . .- . . " . , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

i . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . I

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

..................................................... 2 . . , . " . . . . I .................

-10

v = 87" v = 75"

W = 199000 N

W = I74000 N

(b) Airplane nosedown steps.

Figure 28.- Concluded.

........ . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-10 . .

. . . . . . . . . . . . . . . . . . . I n - - - -1- I k J s - H . . . . . i .. . . . . .. . . . . . . . . . . .

F I . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . . . . . .

.. - .-

VE, knots

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . . . . . .

t '

i f . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1.2 1 1 . . . . . . . . . . . . . . . . . . . . . . .

..............I^.I-.- , ,..-" L - : I : i . . : ..e -..

, . . . . . . . . . . . . . . . ..:: . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

![ f . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

I , . . . . . . . . . . . . . . . . . . . . . . . .

.2 . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

- ...

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

- . . . . . . . . . . . . . . . . . . . . .

.

.I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-

-n , "" -A . . . . . . . . . . . . . . . . .

0 1 " . . - = ? " - . : " " . . . . -

- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-.I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. : . . . . . . . . . . . . . . . . . . . . . . ", . . . . . . . . . . . . . . . . . . . . . . . . . . .

e 1 . .

. . .

v

, . . . . . . . . . . . . . . . . . . . . . . "l\. '

. . . . . . . . . . I . . . " . A : : : : : . : : : . : : : : : . . .

. . , . . " - " .

, . . . . , , , 1 " - . " " " - : - . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 ' .

4 ""-

( - 3 . . -_ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

~ "- . . . . . . . \. ' ; : l ,I. . . .>."A*, . . . . . . -.

Y . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-4 (a) Forward nozzle rotation.

Figure 29.- Response to nozzle rotation, flaps 65".

. . . . . . . . . . . . . . . . . . .

... . . . . . . . ...... ... ... .:.:.... :.. ' 2 , . . . . . . . . . . . . . " . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

F " : f l , i..:.,::,:...;..

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

n -10 . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

- . Y + . - : . - < . r a % " . " . . .

VE, knots

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

1.2 . . . . . . . . . . . . . . . . . . . . .

_ _ ; . ,. .

..-+. . - . . . , , : , * " ..-.;-.. "...- .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

.8 . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

..-.. . . . . . . . . . . . . . . . . . .

! . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . 1

. . . . . . . . . . . . . . . . . . . . -

.2 , ~ ~ - ~ - ~ . , ~ - ~ ~ ~ ~ . - ~ , . ~ ~ "__ ..

~.

, . . ? A " - . . . . . . . . . . . . . . . . . . . .

k -

-. 2

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

j .I . .

. . . . . . . . . . . . . . .

, . . . . . . . . . . . . . . . . . . . .

L i ~ : : . . . . . . . . . . . . .

. . . . . . . . . .

. - *" ...

-

" p..-. . " . . ' . . . . . . . . . . . . . .

-.I , , ; : . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

p . . . . . . . . . . . . . . e . . .

-

I c- .- .%.

"" ?"- ,,-,. ,,A= . . . . . . . ,-- - _I_.

I.. . / ........ ............. ... . . . . . . . . . , . .

. . . . . . . . . . . . . . .

!

. . . . . . . . . . . . . . . . . . . . . . . . . .

I . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.-. . . . . . . . . . . . . . i . . .

. . . . . . . . . . . . . . . . . . . j . . . . . .

i . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

c . , . . . . . . . . . . . . . . . . . i , : !;4 " , . . . . . . . . . . . . . .

"_

i ~-i-

-

.. i . " Be, deg p : : . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . ; ' . . j . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-4 ' . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

~ ....... ,",.~."-~..,-..

~ ~- ~~ ~~ (b) Aft nozzle rotation.

Figure 29.- Continued.

................................ .............. " h., ........................................ I ............

. . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

O t -

-10 I n . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

70 . .:" ..... .( ..... ~ . . . . ~ .~.~ . . ~ . . ~ . . ~ . . ~ . . ~ ~

c

VE, knots I::::::'' ; : ; - . - + +

e / . . . : . . . . . . . .

60 p

. . . . . . . . . . . . . . . . . . . . . . .

1 IO seconds I

. . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

1.2 . . . . . . . . . . . . . . . . . . . . . . . .

8 . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

1 - . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

2 . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

~ "- I

-0-L /"" * .... ~ .___.__

" " , " . < " > . , , < + * _",_ - . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. I I . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . -. . , . . . . . . .

. . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

q, rad/sec 1 :--.----- - ". -

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

-.I . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

[ . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . .

_... . . . . . .

. . . . . . . . . . . .

. . . . . . .

.-.- _..

. . . . . . . .

. . . .

. . . . . . . . . . . . .

I T . : . . . . . . . . . . . . . . . . . , . . . . . . . . . . . . . . . . . . .

r. ... r.. . - , L . . . . . >&* ',.>,,' ... + . '

' J . . . . . . . . . . . . . . . . . . . . . . . .

y . . . . . . . . . . . . . . . . . . . . . .

. . . .

IO0 t-\:,. . . . . : . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

k - - (c) Small nozzle step.

Figure 29.- Concluded.

. * .......... ../ ":"" .......................... L . .......

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . &+.-+-+ """ . . . . . . . I . . . . . . . . . . . . .

h . . " " ' l " " *A

" . .-F-:2*

-10 . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 -

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

................ 3 , ......................... , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

VE, knots . . . . . . . . . . . . . . . . . . . . . . . .

: : : : : : : : : : : : : : . . . . " 4 . . . . . . :?-\:':.:: . . . . . . . . . . . . . -;;?A" w 60 p.. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

: : : : 1 0 seconds : : : : I : : : : : : : : : : : : : : L 4 -I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

- 8 F : : : : : : : : : : : : : : : : : : : : : : : . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

...................................................................................

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

- , - . . . , " " " . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . I "

C " "-

. . . . . . . . . . . . . . . . . . . . . . . .

q radlsec o,lr--.--,--: :+: . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.............................................................. ..............................................

. . . . . . . . . . . . . . . . . . . . . . . .

C Y , : . . . . I . . . . . . . . . . . . . .

. . . . . " ? " < L " * p?. . :.I . " ? . . . .

. . . . . . . . . . . . . - 8 : : ' : : : : . . . . . . . . . . . . . . . . . .

I -I 0 . ._ ....................................................................... ........................... ..............I .............................................

I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

-. . . .

. . . -&./"--..- . . . . . . . . . . . . . .

I . . . . . . . . . . . . . . . . . . . . . . . . I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 0 0 . . . . c _ " 2 "" 1 . ...* " ._ ~~" "" - i~~ - - .""" . , - . . . . . . ..."., .& -...-, i J " 2 _ .................................................................................. .r_ ......................................................................................................................... . . . . .

-..".~.,J.-...% - - """ . . . . . . . . . . . . . . . . . 2.-c " : . . - . . - - . . _...~ . . . . .

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . p . . . . . . . . " : . . . . . . . . . . . . . . . . . . . . . . . NHl . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

"_y___( - . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

O / O rpm

(a) rprn increased, v = 83".

Figure 30.- Response to a throttle step, flaps 65".

8 4 . . .

. . .

& . . . . . . . . . . . . . . . . . . . . . . .

-10 . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . .

. . . . . . . . . . . . . . . . . . . . . . .

. . .

. . . . . . . . . . . . . . . . . . . . . . .

. . .

........................................................................................

. . . . . *"+...&-.~.- *,/"7zr . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

VF, knots

. . . . . . . . . . . . . . . . . . . . . . .

L' " .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

1.2

. . . . . . . . . . . . . . . . . . . . . . . .

.8

.2

C

-. 2

. . . . . . . . . . . . . . . . . . . . . .

C " - .,-~ -5- _ " _i . -

. . . . . . . . . . . . . . . . . . . . . . .

q rad/sec c

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . -:(TF

. . . . . . . . . . . . . . . .

.. .-,'!" " , " . - .. - .,..;,r.i-./- ..... .-

..-<-. .,., ~.

_,., ,..- - ~ . ~ , , / , : , , ' " - " : ~ ' , . . ' , . - . . . . . . . . .

. . . . . . . . . . . .

"~ " .~ " < . " "" --.- . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

A " > "" . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

o / o rpm

(b) rprn decreased, Y = 75'.

Figure 30.- Concluded.

Weight = 189000 to 203000 N

(42600 to 45700 Ib)

NH = 94.3 to 96% rpm

"L-

50 60 70 80 90 1 0 0 1 1 0 120 1 3 0 140 150

Airspeed, VE, knots

Figure 3 1 .- Elevator required for trim for various flap and nozzle deflections.

" . -. "" >,-~ "- -

."". "

c - I 0 seconds-, VE, knots '\

'\

': 600 400 f l "

-1 200

Figure 32.- Transition t o handing configuration, W = 195,000 N.

I '

I I

. . . . . . .

. . .

; " - - I . , . . . . . . . __ -i . , . .- '

NH, % rpm

,P-.. : . : : : : : : . . . : : : : . . . . . .

I(. Right engine ; . . . . . . . . . . . . . . . .

. . . . . . . . . .

-..

. . . . . . . . . . . . . . . . . . . .

" " ... " , . . . . .I . . . . . .

. . . . . . . . . . . . . . .

. . . . . . . .

j

. . . . . . . . . . . . . . . . . . . .

. .I . .

. . . . .

. . . .

1 : . . : " . : . .

. .

"_ _."

. - " - " .: .L" . . .

. . . . . . . . . . .

o r : : : . - . '

-25 . . . . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . . . . . .

. . . . .

. . , . . . .

. . . . . . .

. . .

-

I .I . . . . . . I . , . ' .

. . . .

I .c ........ ?" :.c- *, . .

.- - - -_I .,,. - - ..- . . . . .

. . . . . . . . . . . .

. . . .

. . . . . . . . .

. . . . . .

. . . . . . . . . . . . . . .

I20 seconds

. .

. . . . . . . . .

. . . .

. I 1 : ' : . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . .

I oc

- - . . . .

"_ L . . . . . . . . . . . . . .

. . . . . . . . .

. . . . . .

. . . .

. . . . . . . .

VE knots 5c

Rate of climb, 5

rn /sec 0

. . . .

. . . .

IO0 . . . . . . .

Altitude, rn

. . . . . . . . . .

. . . .

0 ~ . .

(a) Takeoff, 6f = 24.8", W = 21 0,000 N.

Figure 33.- Simulated engine failure time histories.

(a) Concluded.

Figure 33.- Continued.

?

NH, O/O rpm

VE , knots

Rate of climb m /sec Altitude, m (b) Landing approach, W = 165,000 N.

Figure 3 3 .- Con tinued.

I I I I i i I I I . . . .

Left engine . .

- /

I " _

. .

. . .

. . .

/R-ight engine' . . . . . .

I ....... I.

i I ' . . . . . . . . . . .

. .

- .. - . . .

..

- 5 L ' ' , I . . . . . . .

i

20 seconds (b) Concluded.

Figure 33.- Concluded.

. . . . . . . . . . .

. . . . .

. .

. . . . . . . . . . . .

. . . . . . . . . . .

. . . . . .

. . . . . . . .

.......

. . . . . . . . .

. . . . . .

. . . . . . . . . . .

. . . . . . . . . . .

. . . .

. . . . . . . .

. . . . . " " ..........

. . . . " -.

. . . . . .

. . . . . .

. . . . . . .

. . . . .

. . . . . . . .

. . . . . .

. . . . . .

" ........ - . "-

5 !="

."

" -.

" i

.5 t

I O seconds E

" .

_ _ ............. . . .......

............ ." ......................... . . . . . . .

. .

" 25 I - - .

. .

". .....

". . -

-25 ..

- . .- . .

(a) Landing approach, 6 f = 64.7", v = 73", NH = 99.4%, W = 189,000 N.

Figure 34.- Stall time histories.

......

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ., . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .; . . . . . . . . . .

-I . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . I . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. I . . . . . . . . . . . . . . . . 4 . .

"-------;;; . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

- * ? I

. . . . . . . . . . . . .

. . . . . . . . . . . . .

.____ . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

/ ...

- -_.c___ "".~..

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

-10 . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

25) . . _ . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

_.

. . . . . . . . . . . . . . . . . . . . -. . . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . _*. . . . . - .

' ' . * . : ' . . . . . . . ? - r L " .

;I . .-; ..... 1 . .-<. . . . .

. .

" 'V i " ..........

Or . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . .

. . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. < " . . ... "~ -25 . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . , . . . . . . . . . . . .

:j

. . . . . . . . . . .

/----7-;- . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

-5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . .

" - . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . - e + - . . . . . . . . . . . . . . . . . . . . - ~ . . .~ . . . . . . . . . . . . . . . . . .

- .......

" . . . . . . . . . . . . .

K . . . . . . . . . . 7 - . . . . . . . . . . . . . . . . . . "-4 I . . I . " " .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

' 8 ' . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.. - ..____ __ " . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

' , ' ' . _ ' ' . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . >&:->\ . . . . . . . . . . . . . . . .

.-. . . .

y"-"..""-~.-.- 1 .

-

: < . . . . . . . . . . .

. . . . * . . "" .

-

. . . . . . . . . .

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8. . . . . . . . . . . . .

-5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

" .. .~ - __ " . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . # . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . .

. . . . . . . . . . . .

30 I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

. . , . . . . . . . . . . . . . . . . . . . .

20 ' . . . . . . . . . . . .

~ ~ " - - : . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 : : : : : . L . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . " ......

10: . . . . . . . . . . . . . . . . 1 . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . I : : : : : . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .! . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

0: . .I. . 7 .I. . I . .I. . I . .I -.T. .I . .T. . I :I. . I .I. . I . .I. . r: .I. . .T.-TT.l.. I . .I.. I . .I./ I . II.7 (a) Concluded.

Figure 34.- Continued.

I II

. . . . . . . . . . . . . . . . . . . . . .

. . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

" ~ ..

. . . . . . . . . . . . . . . . . .

'Or- . . . . .

. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

- .

. . . . . . . . . . .

5 . - T T . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

C L

' O r "~ ....

5 " . . . . . . . . .

. . . . . . . . . . . .

. .

. . . . . . . . .

. . .

'"e:.

CJ

. . . . . . . . . . .

.5 . .

* . IO seconds , . . .

. . . . . . . . . . . . . . . . . . . . . - - I . . . . . . . . . . . .

" I .5 I" . . . . .

. . .

. . . .

" . . . . . . . . . . . . . .

. .

" - . .

. . . . . .

..

. .

V E ~ knots

""

-

-~ . . . . . . .

. . " . " . " - . . . . . . .

-i

I I . I I I I ~ I ~ t

(b) Landing approach, 6 ~ = 65", v = 74", NH = 96.8%, W = 191,000 N.

Figure 34.- Continued.

" -

i

"" I

IO i '

I

-10 f "

i I

j

. - . ."

.. , - - -- I I . _ 1 I " i i I i ~. ." ".

.-. -.

I -.

~ 1 - 1 0 I

1 . 1 . 1 I. 1 . 1 I I I I l l l l I l 1 - i "

(b) Concluded.

Figure 34.- Continued.

II I

..

5 I : : . . " " . - " _

. . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

"\ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

0 . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . .

. . . .

. . . . .

. .

. . . . . .

a . 'IO seconds-

1.2 .r

- .4L..>.". . . . . . . . . . . . . . . . . . . . . . --

s e

knots

"E 1 .....

"- ".

"."

. . -. , " ! I I I I I I I I l l ~ l l l ! l l ( c ) Takeoff, 6 f = 28.6", v = 1 l o , N H = 9?.6%, W = 192,000 N.

Figure 34.- Continued.

"" ." - .............

(c) Concluded.

34.- Continued.

Figure i, CL CJ -_ (d) Cruise, 6 f = 5.8", v = 7", N H = 90.4%, w = 197,500 N.

Figure 34.- Continued.

-25L j .. - I E+, deg : E . . . . .

-5 I (d) Concluded.

Figure 34.- Concluded.

= 2=3 (CLT = 2.95) I Derivedfromanalysis of

~

b o ( cLT = = ) 1 discrete test conditions

." - . " " " .

Note: Fairings based on parameter identification of

entire data collected in four approaches (CLA=2.65)

(Wing height 1 /(Wing chord)

Figure 35.- Ground effect derived from flight-test data: 6f= 65"; v = 50" t o 80"; N H = 9 3 to 96% rpm; VE = 60 to 70 knots.

. . . . . . . . . . . . . . . . . . . . . . . .

1 . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . r . . . .

- . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

" . . . . "-

I - . .~ . . . - .. - .. -.

n . . . . . . . . . . . . . . . . . . . . . . . . .

- . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . - . . . . . . . . . . . . . . . . . .

-20

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.......................................... ,,-.. ....

. . . . . . . . . . . . . . . . . . . . ~ . . li" L_ _ " . " " " . . I - . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . :A?..; . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . .

. . . . . . . . . . . .

. .

-i.. i , I" . ~- 9 : . . . . . . . . . . . . . . . . . . . . . .

. . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

- . . . . . . . . . . . . . . . . . . . . . . . : : : : 4 seconds : : : : . . . . . . . . . . .

-5 ...............................................................

-

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

-

. . . . . . . . . . . . . . . . . . . . . . . .

.5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ~" - .

. . . . . . . . .

i 7 rad/sec2

. . . . . . . . . . . . . . . . . . . .

- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

; L . ~ I f . : : : : ':-: : - : ; . > .

.2

r , rad/sec

-.2

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

................ . . . . . .....................................

40 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . I . . . . . . . . . . . .

. . . .

~~ . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

................................................................................. .- ...............................................

. . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

p rad/sec2

....................... .............................. - . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

\

p rad/sec

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . :?- "4 i::::::::::::::::::::::: . . . . . . . . . . . . . . : : : : \ : . / : : : : : : : ...................... -.. ~ : : : : : : : : : : : : : ...

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-

50 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ' , . . .

. . . . . . . . . . . . . . . . . .

6, 7 deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . : : : . : : : . . : : : . . . . . . . . . . . . . . . . . . . . . . .%. e . . . .

-

..

. . . . . . . . . . . . . . . . . . . . . . . p . . . . . . : ..<'. . . . . . . . . . . . . . . . . . . . . . .

(a) Flaps 66", 69 knots, SAS off.

Figure 36.- Roll reversal, landing configuration.

I . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. " -,,- __ .."

. . . . . .

. . . . . . . . . . . . . . . .

-25 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . .

. . I . . . . . . . . . _ .............................. -- . . . . . . . . . . . . . . . . .

5 - ~ . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . : : : . . . . . . . . . . . . . . . . . . . . . . . . .

Lateral : : . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . .

. . : : : : / / , . . . : . , ~ , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

SAS servo, " . > r 7 7 7 T X - : \ . " :-= . . . . . / yy -. . . . . . : G ? . T - r . " ; : : ; . . . . . . . . . . . . . . . . . . .

volts

: : : : , . . . . . . . Directional : : : : : : : : : : . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

I . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

-...

-5 . . . . . . . . . . . . . .

. _ .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

J J : : : : : : : : : : : : . : : : : : : : : : : : ] : : : : : : : : . : : : : : : : : : : : : . : ] . . : : : : : : : : : : : : : : : : : : : : : : I . . . . . . . . . . . . . . . . . . . . . . . . . ~ - . .. .L.. . . . . . . .

d- " L " . " . " .... .-. . . .

. . . . . . . . . . . . . . . . . . . . . . 0 . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . - 1 . : : ; . . : : . , : : . . . . . . .

. . . . . . . . . . . . . . . . . . . . , . _ . . . .

-5 . . . . . . . . . . . . . . . . . . . . _ . 4 seconds:: . - - . . . . . . . .

. . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. : . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . .

.

. . . . . . . . . . . . . . . . . . . . . .

" _ __

i , rad/sec2

0 . . . . . . . . . . . . . . . . . . . .

- 5 . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . .

...... . .

. . . . . . . . . . . . . . . . . . . . . .

p rad/sec*

p, rad/sec

(b) Flaps 66", 70 knots, SAS on.

Figure 36.- Concluded.

c

Altitude, NH,

1 x 9 *

kg- m m

rpm,%

C J

I660

. 37 428400 92

. 36 2440 93

366100

I780 0

. 14 340300 89

---Predicted for flight points

- Predicted,sea level standard

NH, O / o Weight, N

178000

I .o

I78000

200000

.2

n

"40 60 80 1 0 0

VE, knots

(a) Flaps 65".

Figure 37.- Lateral control power.

.. . .. .

I I

NH, Altitude, 1x3

m

7 0

C J kg-mz

0 375600 2 130 100

.24

2250 0

34 I700 92

.2 4

.04 2 1.30.

Idle

375600

--- Predicted for flight points (100%

and 92%

rpm)

- Predicted, sea level standard

I .4

Weight, N

I78000

1.2

100 200000

I 1 90 178000

I .o

. 8

. 6

.4

40 60 80 100 120

VE, knots

(b) Flaps 30".

Figure 37.- Continued.

O L

100 120 140 160 180

VE, knots

(c) Flaps 5.6".

Figure 37.- Concluded.

Ailerons,chokes,and spoilers

Shaded - SAS o f f

. 7

Open - SAS on

I 9 A / I

I I I

. 6

U

off

L * /

n A - . A 2

- .4

.a

A&o-A m -

A-

off

Chokes and spoilers

0 20 40 60 80

Control wheelangle, &, deg

(a) Effectiveness of combined lateral controls.

Figure 38.- Lateral control effectiveness, flaps 67", airspeed 69 knots, 15" nozzle, W = 165,000 to 178,000 N, NH = 93 to 95% rpm.

.3

I !

20 40 60

Wheel deflection, Bw, deg

(b) Effectiveness of individual lateral controls.

Figure 38.- Concluded.

NH

7 0

0 93-95

0 89

Flagged - chokes inoperative

Shaded - SAS on

7 Open - SAS off \ H N =93-950/0 rpm

c I

. 4

. 3

.2

.I

0 20 40 60 80

Wheel angle, SW, deg

Figure 39.- Effect of engine power and SAS on lateral control effectiveness, flaps 67", 69 knots, v = 15O, W = 165,000 to 178,000 N.

. 6

. 4

.2

20 40 60

Wheel angle, 6,, deg

NH9 "E9

O/O

knots

-

96.5

62 -46

o m

93 69 -35

A

89 69

.I4

0 93 72

.34

-

h

95 71-77

0 69 .35

d

96 72 .34

Shaded- SAS off

Open- SAS on

I .ox --

0.8

0.6

0.4

0.2

0 20 40 60 80 I O 0

Wheel deflection, S W , deg

Figure 41 .- Lateral control effectiveness, flaps 65", W = 168,000 to 182,000 N.

Weight,

70 199000

/

I O 20 30 40 50 60 70 80 90

Wheel angle, 8w, deg

Figure 42.- Lateral control effectiveness, landing flaps, 69 knots, N H = 95%, SAS on.

1 1 1

v = 75 deg

Weight = 176000 N

NH = idle (Left-hand engine)

= 99.7% (Right-hand engine)

NH

- . 6

-80 -60 -40 -20 0 20 40 60 80

Wheel deflection, &, deg

Figure 43.- Lateral control effectiveness for simulated one engine inoperative, flaps 66", 72 knots.

u = 10"

Weight= 165000 to 182000 N

FI agged symbol - control removed

before max rate reached p 4 r ! ! ' o f f

20 40 60 80 100

Wheel angle, sw, deg

Figure 44.- Maximum roll rate from lateral control reversals, flaps 65", 69 knots, W = 165,000 to 182,000 N, nozzle deflection = 10'.

Weight, v, NH, 4 ,

O/O knots SAS N

deg deg

~

1 5 92 7% on 200000

0 33

0 33 1 5 92 7% I98000

off

0 5.6 7 92 Off 200000

I96000

n 5.6 7 90 I34 Off

20 30 40 50 60 70 80

0 IO

Wheel angle, Bw, deg

Figure 45.- Lateral control effectiveness, flaps 33" and 5.6".

I I 1 N H V Weight,

O/O N

0 0

200000

. O I98000 92

e0 200000 92

Shaded - SAS off

Open - SAS on

Unflagged- clockwise s w

Flagged - counterclockwise SW

-

I I I I

20 40 60 80

Wheel angle, sw, deg

Figure 46.- Lateral control effectiveness, flaps -30°, 78 knots.

- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

?..:.. . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

........................

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

- . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

-40

*-.-..""."." .................. . . . . I . . . . . . 1 ...........

............. ................... " ............................... *" . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

~ ~____ . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

-20

-..- -."_....-..""_.. "" - ................... .........."........

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.4

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . :p.,; : . . . . . . . . . . . . . . . . . . . . . . . . " / X ; . . . . . .\

radlsec 2 0

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .

- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-. 4

""".-...-....-.- - .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.2

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

- "

" " - ""c

rad/sec

. . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-.2

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

................................................................

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.4

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . : : : : : : : / - \ : . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . , . . . . . .

,' / ~~ ~ " x

0 1 - \

. . . . . . . . . . . . . . . . . . /. -*-; .. -. . . . . . . . .

rad/sec2

. . . . . . . . . . . . . . . . . . ) . " . . . . . . . . . . . . . .

?L- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-.4

- - ".. ..... -.- " .- ................ .................................................................... " . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.2

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

--.

"

. . . . "--- ___

rad/sec

i s . . . . . . . . . .

-. 2

. . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . .

- 20

I O

. . . . . . . . . . . . . .

- . 7+=

. . . . . . . . . . . . .

. . . . . . r: . . . .

. . . . . . . ,w. : . . . .

. . . . . . . . . . . . . . . . .

I

-10

Figure 47.- Rudder reversal - flaps 67", 68 knots, SAS off.

... ".......,.............I.. ........................................................

...................................................... .........................

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . .

. l i . . . . . . . . . . . . . . . . . . .

:: : : ://. Direction

. . . . . . . . . . . . . . /

SAS serva, ......................... 7 "===Z . . . . . . . . c . .

. 4.: volts . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . ' Lateral

..... ".."__.."..-...1..1. ......... .."... ............ ... * ................................... ..r ^ .._.......-.......

-5

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . :/ Right : . . : . . . . . . . . . . . . . . . . . . . . . I

-

?"x-....._::- . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

A " . ~ c

-

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.2

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

*h " ." " *

. . x. T : . : .-. . :?r+

. . . . . . . . . . . . . . . . . . . . . . . .

p rad/sec2

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-. 2

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

_.. ....".... - ....-...............__ "............................. ........ ............

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

IO . . . . . . . . . . . . . . . . . . . . . . .

"?-. - . , - ~- x . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-10

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.......... "... _.. ........................................................... ............................................................................................

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.4

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . y::: a , . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

" - 1

- ' i

rad/sec2

0 "\

r , . . . . . . . . . . . . . . . . . . . . . . .I,. . . . . ./. . .

. . . . . . . . . + . + f : : : : : : : . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

-. 4

...............................................................................

.......................... -. ....................................................

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

.2

. . . . . . . . . . . . . . . . . . . . . . . . . /?\. . . . . . . . . .

-

rad/sec 0

. . . . . . . . . . . . . d . . . . . . . . . . . . . " r, . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . .

-. 2

. . . . . . . . . . . . . . . . . . . . . . . . .

,".,*, - . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

-20

I O

.c- '. . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

/ . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . .

-10

Figure 48.- Rudder reversal - flaps 67", 70 knots, SAS on.

o 70 knots, = 66" I / = 8"

Iz= 6 27,000 kg-rn2

0 72.5 knots, = 65", V = I O "

Iz=631,000 kg -m2

/

-Predicted 72.5 knots

Figure 49.- Rudder effectiveness (after empennage modification).

Predicted,

.3

67"

1 5 "

Nu .2

%

\ VE= 68 to 7 1 knots, D

.I

Weight = 192000N

. L !

.4

/ - y B f = 33"

.3

v = 7" to I 4 O

O J W

? .2

0 VE = 78 to 82 knots, U

E

Weight = 193000N

L a

.I

0 VE = 76 to 78 knots,

Weight = 180000N

0 4 8 1 2 1 6 20 24

ABrl deg Figure 50.- Rudder effectiveness (before empennage modification).

d

N knots

deg deg

I91 000 69 15 67

1 0

Flags:Controlremovedbefore max rate reached

a Q) t tl L

x

0 5 IO 1 5 20 25 0 5 IO

&, deg, measured from - I o trim 8 p 9 cm

Figure 5 1 .- Maximum yaw rate from directional control reversals (before empennage modification).

IO

- 5

-10

- I 5

J Elevator angle

Be, deg O(

- 5

-10

- I 5

- I 5 - I 0 -5 0 5 1 0 1 5

P, deg

Figure 52.- Steady sideslip - 6 f = 65", v = 15", VE = 65 knots, NH = 93%.

0 Before empennage modification 0 After empennage modification

9-

I O

0 5

a ,

O O

~ -5

I

-10

I

-I 5

I

- 20

I I I Predicted I

IO

0 5

W U

s

60 - 5

I

-10

I

-I 5

-20

I O

-& -5

-10

-I 5

Figure 53.- Steady sideslip, flaps 65", VE = 65 knots, v = 15" , NH = 93%.

"E

A 97 knots, after empennage modification

013 65 knots

} before empennage modification

0 90 knots

L

" $-IO

P - 2 0

-3 0

Y -10 rn

-I 5

P, deg

Figure 54.- Steady sideslip, flaps 30°, v = 6", NH = 90%.

, I

I O Wheel, rudder angle

-I 0

9 deg

-20

-30

I O - 5

-I 0

- I 5 I O Elevator angle Be, deg

- 5

-10

- I5 -10 - 5 0 5 I O 1 5 20

rs, deg

Figure 55.- Steady sideslip - 6 f = 5.6", V r = 120 knots.

Y J

I O

Wheel, rudder

angle 0

B w , B r , deg -10

-20

-30

I O

- 5

-10

-15

I O

Elevator angle Be, deg

-5

-10

- 5 0 5 I O 15

- I 5 -10

P, deg

Figure 56.- Steady sideslip - 6 f = 5.6", VE = 1 5 0 knots.

Weight, N

deg ~- deg I SAS

67 - 1 0 I65000 to I82000 Off 67 57-86 165000 to 182000 off 1 7 I72000 on 62 82 194000

on *

1 5 202000 off A

33 9

I76000 off

*

1 5 195000

on

*

*

30 II 200000 on

n

6 I1 I96000 O f f

n I 6

I 1 I76000 off

* Flagged- SAS modified

Shaded - SAS on