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https://ntrs.nasa.gov/search.jsp?R=19800023830 2018-06-17T17:01:00+00:00Z

B, Feinreich and G. GevaertLear Siegler, Inc,Astronics Division, Santa Monica, CA

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NASA CR 152365

A COMPARISON OF FLIGHT AND SIMULATION DATA FOR THREE

AUTOMATIC LANDING SYSTEM CONTROL LAWS FOR THE

AUGMENTOR WING JET STOL RESEARCH AIRPLANE

(NASA-CH- 152365) A COdPAI iSON OF FLIGHT AND S60-32338SIMULATION DATA FOU THREE AUTOMATIC LANDINuSYSTEM CONTROL LAWS FOR VIP, AUOMENTOR WINOJET STOL UBSEARCH AI0LAhE (Lear Siegler, linclal;

Inc.) 18 p [IC. .A02IMF A01 C,SCL 01A X3101 34110

Abstract

The extensive technology base whichexists for automatic flare and decrabcontrol laws for conventional takeoffand landing aircraft has been adaptedto the unique requirements of thepowered lift STOL airplane. Threelongitudinal autoland control laws weredeveloped. In addition to conventionalcontrollers, direct lift and directdrag control were used in thelongitudinal axis. A fast timesimulation was used for the control lawsynthesis, with emphasis on stochasticperformance prediction and evaluation.Through iterative refinements, goodcorrelation with flight test resultswas obtained. This simulation was usedto extrapolate the statistical landingdata base beyond the two sigma levelestablished in flight to the improbablelevel required by the FAA forcertification. Excellent touchdownsink-rate control was obtained, withrange accuracy consistent with Cat IIIperformance requirements.

Introduction

The Ames Research Center of NASA isconducting a series of investigationsto generate and verify through groundbased simulation and flight research adata base to aid in the design andcertification of advanced propulsivelift short takeoff and landing (STOL)aircraft. One portion of this programis concerned with obtaining technicalinformation on automatic landingsystems for STOL aircraft includingflight path control performance andtouchdown state dispersion in thepresence of environmentaldistrurbances. As part of thisprogram, Lear Siegler's AstronicsDivision developed automatic landingcontrol laws for the Augmentor Wing JetSTOL Research Airplane.

The technology for the development andcertification of Category III automaticlanding systems for conventionaltakeoff and landing (CTOL) jettransports is well developed anddocumented, as noted in references 1 to3 for one commercial aircraft andreference 4 for FAA requirements. Nocomparable technology exists forautomatic landing systems for STOLairplanes in general and for poweredlift STOL airplanes in particular.

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The objective of the automatic landingwork reported here is to gainunderstanding of the problems impactingthe design of powered lift short-haulairplanes that are to be landedautomatically on STOL runways inadverse weather conditions. Thisunderstanding was attained by a limitedcoverage of'important elements that arenormally included in the certificationprocess of a CAT III automatic landingsystem for CTOL airplanes with majoremphasis on fault-free performance.The control law developmentconcentrated on the final approach to

touchdown phase of the 1-:ndi'ng, withthe majority of the effov l t expended onlongitudinal and vertical Controlbecause this is where the peculiaritiesof the powered-lift STOL vehicle aremost prominent.

The development and flight validationof the automatic landing system controllaws was conducted in three phases. Inthe first phase, reported in reference5, Lear Siegler developed bothlongitudinal and lateral candidateautoland control laws for a poweredlift STOL airplane using an AugmentorWing Jet STOL airplane as an example.This development was based on previouscompany experience with automaticlanding system designs and on controlstrategies which were emerging frommanual operation of the Augmentor Wingairplane by NASA pilots. Fordiscussion of these manual operations,see reference 6.

In phase 2, candidate automatic landingcontrol laws were selected by NASA forimplementation. NASA personnelsupervised the development andqualification of the flight software onthe airborne hardware simulationresident at the Ames Research Center,conducted the flight testing andanalyzed the performance of thesecontrol laws.

As the flight program progressed,models and control laws were refined ina Joint effort of Lear-Siegler andNASA, culminating in the configurationspresented in this report.

Although a lateral control law wasflight qualified and evaluated, themain thrust of the program remained onthe longitudinal control laws. Threelongitudinal control laws emerged forcomparison. The primary emphasis inall three longitudinal laws was onachieving an accurate touchdown sink-rate with secondary emphasis on touch-down range dispersion.

In the third phase of the program, LearSiegler used the results of the NASAflight testing to validate a high speedanalog simulation which was then usedto generate a large statistical database to establish the automatic landingsystem performance at the 10-6probability (improbable event) level.

This paper describes the development ofa family of automatic landing systemcontrol laws and shows that this typeof control law is capable of meetingrequirements like those applied by theFAA to CTOL automatic landing systems.The results presented in this paper arederived from both simulation and flightdata. A comparison of flight and simu-lation establishes the validity of thesimulation both as a design tool and asa mechanism for extrapolating theflight data to the improbable event.

The Research Airplane and the ApproachoLnTi' tons

The Airplane

The Augmentor Wing airplane shown inFigure 1 is a modified de Havill•andC-8A Buffalo airplane with the wingspan reduced to increase wing loading.

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This airplane is equipped with jet aug-mentor flaps as shown in Figure 2, in-corporating flow blocking devicescalled chokes, has drooped aileronswith boundary layer control and in-corporates full span leading edgeslats. The two original turbopropengines were replaced by two RollsRoyce Spey 801 split flow turbo-fanengines which were supplied by theCanadian government as part of thejoint program between NASA AmesResearch Center and the CanadianDepartment of Trade, Industry andCommerce. The cold flow from theengines is ducted to the augmentorflaps and ailerons and the hot thrustis vectorable through the conicalexhaust nozzles. A more detaileddescription of the aircraft and itscharacteristics is given in Reference 7.

The Approach Condition

The nominal landing approach conditionsare given in Table I.

Table I Nominal Approach Conditions

Airspeed, knots 70Gross Wight, lbs 43,000Wing Loading lb/ft 2 49.7Lift Coefficient 3.0Flight Path Angle, degrees -7.5Augmentor Wing Flaps

Deflection, degrees 65Engine RPM, percent 95Thrust Diverter Nozzles

Deflection, degreesfrom horizontal 75

The Augmentor Wing Jet STOL Researchairplane was flown on a 7.5 degreeglideslope at speeds near 70 knots forthe final approach. At this lowapproach speed the airplane operates onthe backside of the power curve.Because of this and the near verticalthrust orientation in the approachconfiguration the most effectivecontrol for path is the throttle andthe most effective control for speed isthe elevator.

These characteristics are in sharpcontrast to a conventional jet trans-port where during the approach the oathis primarily controlled with the ele-vator and the speed is primarilycontrolled with the throttle. Refer-ence 6 contains a more complete Jis-cussion of the operating character-istics of the Augmentor Wing airplane.

The Airplane Controls

The Augmentor Wing airplane incor-porates four controls that can be usedin the longitudinal axis for thecontrol of glidepath and automaticflare. The throttle regulates RPMwhich in turn regulates hot thrustthrough the exhaust nozzles and coldthrust through the augmentor flaps.The autothrottle was mechanized to givea lift control authority of +0.1a and-0.07g's about the nominal trim pointwhile observing engine limitations andpreserving lift margins. Direct liftcontrol is available through the sym.metric actuation of surfaces called,^hQkes which can block the flow throughthe inboard augmentor flaps. Thesefast acting chokes, when used, aremodulated k30 percent of full closure

about a nominal 30 percent position toprovide approximately 10.1q's of liftauthority. When the chokes are used,they are complemented with the thrnttloto improve overall path control hand-width at the expense of some overallreduction in powered lift augmen-tation. The powered lift lost by

biasing the chokes must be re p laced byincreasing the aerodynamic lift througha small increase in approach reference

airspeed. The thrust conical nozzles,which can be vectored from 6 0 to 1040from horizontal are always used to trim

engine RPM and for some control config-urations are also used as a direct

longitudinal force device for shortterm speed control. As a trim device,the nozzles are adjusted to compensatefor temperature and wind in order tomaintain the engine RPM in a nominal

operating range to provide: for bothupward and downward path corrections.

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The maximum RPM limitation is estab-

lished to avoid structural damage tothe nozzles when the nozzles are down.The minimum RPM is set to maintain aminimum value of lift margin as des -cribed subsequently. When used as alongitudinal speed control device, thenozzles have a longitudinal authorityof +0.13 and -0.09g's for typicalnozzle trim values near 75% A hy-draulic powered elevator is the fourthcontrol which is always used for longterm speed corrections and is alsoused, in the absence of nozzlevector { ng, to provide short term speedcontrol.

Roll is controlled with ailerons,spoilers and outboard augmentor flapchokes which are mechanically geared tothe wheel. A split segment but other-wise conventional powered rudder isused to control yaw.

A unique characteristic of a poweredlift aircraft is that it can approachat speeds below the power off stallspeed. In order to provide adequatesafety margins, CTOL aircraft use anapproach speed of 1.3 times the poweroff stall speed. For powered liftaircraft this would be an excessiverequirement and other means must beused to provide safety marginscomparbie to that used for the CTOLvehicle. Reference 8 describes acomprehensive study of this problem.On the Augmentor Wing aircraft, a liftmargin of 0.4g ensures a safe approachspeed. Lift margin is defined as thedifference in g's between the trim liftvalue and the maximum lift availablefrom pitch rotation alone with thethrottle held constant. Since the liftmargin is a function of sped andthrust, limits must be placed not onlyon the approach speed but also on theminimum value of engine RPM.

The Avionics System

The airplane is equipped with theSTOLAND digital avionics system(Reference 9) providing versatilenavigation, guidance, control anddisplay functions.

A microwave landing system was used forapproach guidance, providing azimuth,elevation and distance information.

Design and Evaluation Methods

Design and Evaluation Process

The design and evaluation process usedin this program includes several of themajor elements that constitute thecertification process of a CTOLairplane CAT III automatic landingsystem as reported in references 1through 3. Figure 3 depicts the majorelements and flow paths included in thecurrent program. A simulation is usedto define and refine the control lawsand verify that they produce acceptablelanding performance with environmentaldistrurbances. Initial flight testresults are used to refine control lawsand airframe models used in thesimulation. When good correlation isestablished between flight andsimulation results, the simulation canbe used to expand the limitedstatistical flight data base (=102)to the extreme event levels (=106)required for certification.

Using the simulation, data was takenfor various levels of environmentaldistuubances, airframe variations andsystem errors, covering a wider rangethan possible in flight. Probabilitydistributions were generated for alltouchdown state variables. Totalpopulation probability distributionswere obtained by convolving thedistributions contributed by thedifferent disturbances.

In two major areas this program wasless comprehensive than a fullcertification program; Heavy emphasisduring the control law development wasplaced on performance with no systemfailures. Less consideration was givento failure effects and redundancyrequirements. The system flown wasnonredundant, relying on pilotmonitoring to ensure safety.

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In a full certification program,correlation between simulation andflight is verified through thecollection of actual disturbance dataas encountered in flight on a landingby landing basis, inserting the samedisturbances in the simulation, andcorrelating the results for a limitednumber of landings. This was not donein this program due contract fundingconstraints. Total population resultsfor a given control law configurationis used for correlation instead.

Simulation

A fast-time simulation was the majortool used in synthesizing andevaluating the automatic landingcontrol laws. Mathematical models ofthe airframe, controllers, sensors andthe environment were assembled and usedin the simulation. The normal set ofuncoupled, linearized, smallperturbation equations of motion wereused in separate longitudinal andlateral simulations. Longitudinaldynamics were included in the lateralsimulation to the extent necessary toaccount for the goundspeeds associatedwith different headwinds. Importantnonlinearities were modeled, includinglift and drag variations associatedwith changing nozzle angles (which arenot small) and with engine RPMsettings; lift, pitching moment, anddrag variation due to ground effects.

Controller dynamics were modeled,'--.including rate and position limits and,significant hysteresis effects.Special care was taken in accuratelymodelling engine dynamic. iecause theengine is used as the major flight pathangle controller and has a strongimpact on performance. Enginemodelling was based on theidentification work described inPeference 10. Separate paths were usedfor computing cold and hot thrustresponses, with different timeconstants used for thrust increase ordecrease.

Sensor dynamics and error models whichcontribute to landing dispersions werealso included, such as radar altimeterdynamics and offsets, and dynamic andstatic vertical gyro and accelerometererrors. MLS noise was modeled andincluded in the simulation. Minds,shears and turbulence consistent withthe definitions in the FAA AdvisoryCircular 20-57A (reference 4) were used.

For statistical dat .r collection, thesimulation was run in fast timerepetitive operation mode, startin q at1000 feet above the runway with theairplane stabilized on the glide-slopeor localizer, and terminating attouchdown. The 100 foot approachwindow states were recorded, as werethe touchdown states: vertical andlateral velocity, touchdown point onthe runway, and pitch, roll and headingangles.

Hardware Simulation and Flight Test s

The automatic landing control lairs wereprogrammed into flight control computersoftware, with testing and validationon the ,JASA Ames Research Center realtime hardware simulator. This totalnonlinear six degrees of freedomsimulation includes flight control anddisplay computer hardware and pilotinterface. The simulation facility wasused to q ualify each software revisionprior to flight.

Flight tests were conducted by NASAAmes Research Center at Crows LandingNaval Auxillary Landing Field (NALF) inCalifornia. The flight test landingswere made on a simulated 1700 by 100feet STOL runway with boundariespainted, in accordance with reference11, on a longer and wider runway. Amicrowave landing system wasinstalled. A data collection andreduction system with airborne andground based elements was used torecord flight test results.

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Control Laws Description

Longitudinal

The glide slope track and flare controllaws that have been developed for theAugmentor Wing airplane are shown inthe block diagram of Figure 4. A backside of the power curve controltechnique is used, controlling flightpath angle with'engine RPM, augmentedby the DLC chokes. The elevator isused for attitude stabilization andcontrol and for long term airspeed trimchanges. Short term airspeeddeviations are controlled tf;rourjh theuse of the conical nozzles which arcalso used for longitudinal trim controlto account for the aerodynamic flightpath angles resulting from differingwind components. The trim tables shownin Figure 4 pre-position the throttles,nozzles and pitch attitude, and theclosed loop control laws correct fordeviations from trim. The trim tablesoutputs are held constant below 300feet radar altitude. Raw glide slopedeviation, computed from elevation andrange information, is combined withvertical acceleration in acomplementary filter to produceestimates of glide slope deviation andrate which are used for tracking theglide slope. The output of the radaraltimeter is blended with verticalacceleration in another complementaryfilter to produce a sink rate signalthat is used in the flare. The glideslope error is faded out prior to flareinitiation. Through the flare, derivedsink rate is transitioned from glideslope to radar altimeter basedinformation, minimizing the impact ofterrain irregularities. A straightline h/ti profile from the existingpre-flare sink rate to the desiredtouchdown value is commanded in theflare.

Vertical path errors generate athrottle position or normalacceleration command which drivesengine RPM and DLC chokes in acomplementary combination.

Engine RPM and throttle position areused as feedbacks for the throttle loopto quicken engine response and minimizethe effects of hysteresis in thethrottle cables. A lag of about onesecond is associated with theunaugmented engine RPM response to athrottle position change. The closedloop response of the throttle servo andengine to throttle position command canbe approximated by second orderdynamics. In flight, with a properchoice of gains, a natural frequency ofup to 2.5 radians per second(critically damped) could be obtained;attempts to further increase thebandwidth resulted in ringing primarilydue to the low rate capability of thethrottle servo which was designed forCTOL applications. The chokes aredriven with the error between throttleposition command and engine RPM,providing fast normal accelerationwhile engine response is building up.Variations in throttle bandwidth wereused to obtain a desirable mix betweenengine and choke activity whileretaining the same overall bandwidth.

Pitch attitude and rate feedback to theelevator are used in stabilizingattitude. On the glide slope, thepitch attitude command provides longterm speed control by summingintegrated raw airspeed error with thetrim table output. Through the flare,attitude is ramped with decreasingaltitude from its pre-flare value tothe desired touchdown value. Thishelps arrest the sink rate and puts theairplane in a proper attitude fortouchdown. This form of control law issimilar to the technique used by pilotsfor manual landings of the AuqmentorWing Airplane and manual landing ofCTOL aircraft. Pitch rotation startsat a main gear height of 65 feetwhereas sink rate flare command startsat 50 feet. The proper phasing betweenthese two events provides a smoothentry into the flare.

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Raw airspeed is blended in acomplementary filter with longitudinalacceleration to produce an estimate ofairspeed error which drives thediverter nozzles. A decelerationcommand is applied during the flare inorder to touchdown at approximately 60

knots.

The control-laws described aboveutilize all four controllers availablein pitch; configurations using threeand two controllers were also definedand evaluated in flight. This was donein order to establish the tradeoffbetween landing accuracy obtainable by

using all controllers and systemsimplicity gained by minimizing thenumber'of active controllers. Table IIsummarizes the allocation ofcontrollers in the different controllaw configurations.

Table II Controller Allocation

Controllers: 4 3 2Throttle it Y YElevator a V,e V,eChoke y y -Nozzle V - -

The nozzle is used for longitudinaltrim control on all configurations.All three control law configurationsare shown in Figure 4. For the four-control configuration, Kh and KVgare zero and K,-y and KVN arenonzero. For the three-controlconfiguration, KK and Kvn are zeroand Kc h and KV are non-zero. Forthe two-control configuration, KCNand KVN are zero and Kh and KVGare non-zero.

Lateral-Directional

Figure 5 is a block diagram of thelocalizer track and runway alignmentcontrol laws. Roll control on theAugmentor Wing airplane is achieved bymechanically linki;ig the aileron, rollspoiler and outboard chokes to thecontrol wheel. The lateral control lawoutput commands a wheel position forroll control. Raw localizer lateral

displacement, computed from azimuthangle deviation And range, is blena dwith cross track acceleration in acomplementary filter. Yaw accelerationis also added as an input to the filterin order to convert lateralacceleration at the center of gravityto the value at the localizer antenna,located at the airplane's nose. Theestimated localizer deviation and itsrate are used to command bank angle.The yaw rate, lateral acceleration andbank angle signals are fed throughgains, summed and gain scheduled withdynamic pressure to drive the rudderfor yaw stability augmentation and turncoordination.

A forward slip maneuver is used forrunway alignment. Beginning at analtitude of 150 feet, an align coixnandis switched into the yaw axis. Thisreference heading command is reducedfrom the heading error existing atalignment initiation to zero at 50feet, yielding an alignment rate whichis a function of both initial headingerror and aircraft sink rate. Theerror from the commanded headingtrajectory is integrated to maintainthe steady rudder required duringalignment. In the roll axis, the beamcomputations are maintained to guidethe vehicle along the desiredhorizontal path, with increased crosstrack rate gain for better control. {abank command proportional to laggedlateral acceleration is added in alignto compensate for sideslip inducedcross track acceleration. A rollkicker is switched in at align toprovide a predictive bank command basedon initial heading error. Bankcommands in the localizer track pathand in the align path are limited to*100 and *50 respectively, which isample authority to handle steady crosswinds to excess of 15 knots.

Landing Performance R esults

Longitudinal.

Figures 6 through 11 show touchdownsink rate and range proba0 litydistributions obtained with thesimulation for the four, three and t-wocontrol configurations with flight-testdata points superimposed.

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The appendix; explains how to read thesecurves for the benefit of the readerswho are not familiar with this form ofdata presentation. The simulationresults were taken with limiting windsand shears, limiting turbulence and MLSnoise. The curves shown are based on a70 percent probability of encounteringlimiting headwnds and 30 percentprobability for limiting tailwinds(limiting headwi nds have a magnitude of25 knots and limiting tailwinds are 10knots). Flight results are based on 31landings with the four controlconfiguration, 29 landings with threecontrols and 26 two control landings.

A fairly wide range of ambientconditions were encountered during theflight tests since the flights wereconducted over several months of theyear and different hours of the day.The distribution of winds measured at amast near the touchdown zone is shownin figure 12 for the four controlconfiguration tests. Even though themajority of landings were made in lightwinds, headwi nds of up to 15 knots,tailwinds up to 11 knots and crosswindsup to 20 knots were encountered. Inthis program, correlation of flighttest and simulator results on a landingby landing basis was not attempted.

Overall probability distributionsobtained in flight are compared withthe simulator generated probabilitydistributions. These probabilitydistributions are best compared interms of their slopes. When makingthis comparison, steeper slopes areexpected in the flight data becuaseflying occurred in less than limitingwind conditions. Weight andtemperature variations and sensorerrors (such as radar altimeter bias)in flight tend to decrease theprobability distribution curve slopeand reduce the difference betweenflight and simulation.

The simulator curve of Figure 6indicates that excellent sink ratecontrol was produced by the fourcontrol configuration.

The predicted mean touchdown sink rateis 3.8 fps, the two sigma hard landingsink rate is 5.5 fps and the sink rateat a probability level of 10•--6 is 7.6fps, which is well within the allowablemaximum sink rate of 12 fps for theAugmentor Wing airplane. Theprobability distribution slope of theflight test points is somewhat steeperthan that obtained in the simulationand this trend is expected as discussedabove.

Figure 7 shows the touchdown sink rateresults for the three controlconfiguration. Since the verticalchannel of the four and three controlsystems are identical, the performanceof the two systems should be similar.Figures 6 and 7 show that thesimulation data for the two systems isessentially the same. The flight datafor the three control system also has asteeper slope than the simulationdata. *

Si nkrate distribution for the twocontrol system is shown in Figure S.Here the highest tolerable bandwidth inthe throttle loop was used but thepredicted sink rate probabilitydistribution curve slope is =flatterthan with the four and three controlsystems becuase of the reduction inbandwidth associated with this no chokeconfiguration. Mere again, the slopein the sink rate probabilitydistribution curve was steeper inflight than in the simulation.

* After collecting the three controlflight data, an error in implementationof the engine: - choke system wasdiscovered. Since the error increasedthe effective choke gain and reducedthe effective engine gain the overallnormal acceleration was unchanged aslong as the chokes were not driven totheir limits. This was the case forthe flight data shown in Figure 7. Asubsequent simulator check alsoconfirmed that for the disturbancesexperienced in flight, the existingflight data was valid.

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The data point at 6 fps tends to followthe bend in the distribution curvepredicted by the simulation.

The touchdown sink rare performance ofthe three longitudinal control laws issummarized :in Table M. The perfor-mance of all three control laws issatisfactory in the two sigma area.Good agreement exists between theflight and simulation touchdown sinkrate results. The 10-6 sink rateperformance.f alts within the 12 fpscapability of the airplane for allthree configurations. However, the10-6 sink rate performance for thefour and three control system is con-siderably better than that of the twocontrol system.

Figures 9. 10, 11 show the touchdownrange distribution results for the samethree configurations. Range isreferenced to the glide path interceptpoint (GPIP). GPIP is 80 feet short ofthe touchdown zone as defined inReference 11 and painted on the CrowsLanding NALF STOL port. Two simulatorcurves are given in each figure, oneshowing the probability of landing longand the other for landing short.

For the four controls, the simulationdetermined mean touchdown point is 310feet with 400 feet between two sigmaland short and two sigma land long asshown in Figure 9. Most of the flightpoints are in good agreement with thesimulation results. This is even truefor the top four points which are asso-ciated with headwinds in excess of 15knots, crosswinds between 15 and 20knots and approaches that were not wellstabilized. The crosswind conditionswere beyond the design envelope of thesystem.

The simulation data in Figure 10 forthe three control system is essentiallythe same as the simulation data for thefour control system. Again, the flightdata correlate well with the simulationdata and the flight data shows theexpected steeper slope associated withlighter winds. The mean touchdown dis-tance for the three control flight datais 70 feet longer than the meandistance for the four control data.

The simulation data for the two controlsystem shown in Figure 11 predictsnearly the same mean range but a 50percent increase in the short landingto long landing range spread as com-pared to the four and three controlsytems.

Table Ill Touchdown Performance Comparison

Variable

Four Control Three Control Two Control

Sink rate, fps ----mean 3.8 3.8 3.1 3.7 3.3 3.7two sigma, hard 5.0 5.5 4.3 5.2 5.6 6.210-6 7.6 7.4 11.9

Range ftmean 250 310 320 300 170 290two sigma, short -30 100 150 100 -95 30two sigma, long 380 500 440 520 400 65010-6 short -210 -160 -32010-6 long 760 780 1040Comments: 1) Simulation results are with limiting winds and shears,

limiting turbulence and ML3-beam noise.2) Range is measured from the Glidepath Intercept Point.

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Thi slope of the flight dataProbability curve is in reasonably goodagreement with the simulation resultsbut the mean value of the flight datais 120 feet short. Differences in themean touchdown range values betweenflight and simulation are probably theresults of a residual modelingdiscrepance coupled with the fact thatrange is not explicitly controlled.

The touchdown range performancecomparison of the three logitudinalcontrol laws is shown in Table III,Good agreement exists between flightand simulation touchdown range resultswith the exception of the difference inthe mean value.

The data contained in Table IIIincludes the kind of performancenumbers that are re quired for CTOIautoland certification. The touchdownranges shown constitute a largepercentage of the 1500 to 1800 footSTOL runway length called for in theplanning document for ST01, .)orts(reference 11). Certainl, when thereis a premium on redncing the touchdownrange dispersion, as is the case forthe STOL airplane, the betterperformance of the four and threecontrol systems is to be preferred.Improvement in touchdown range controlmay be obtained by commanding a highertouchdown sink rate.

In addition to sink rate and range,pitch attitude is a touchdown parameterof interest. Pitch attitude should behigh enough such that the airplanewould land on themain wheels prior toallowing the nose wheel to touch therunway but it should not be so high asto allow contact of the lower aftfuselage with the runway. Touchdownpitch attitude was well controlled forthe Augmentor Wing Airplane for allthree configurations. A F' mean wasobtained with about 1' two sigmadispersion. At the 10-6 level, thepitch attitude is well within the -1,+15 degree boundaries determined fromthis airplane's geometry.

Lateral/ (ir choralFigure 13 shnws thr* latr:r,al touchdowndistance distribution obtained with thesimulation 4nd in 67 landings inM ght. The simulation data wa y ta. enwith maximum design (15 knots)crosswinds and turbulence. The spreadin the lateral touchdown distributionof the flight data is more than doublethat obtained from the simulation. Theextreme deviations to the right of therunway's centerline (beyond 15 feet),shown by flight data, are associated66th the system operating near orbeyond its limits, with quarteringheadwinds of more than 20 knots and aleft crosswind component up to 20knots, resulting in rudder limiting insome cases, This, however, does notexplain the overall wider lateraltouchdown distribution of the flightdata which is a result of a problemthat has not been pinpointed. Othermanifestations of this problem are rollexcursions from side to side duringalignment and not a very tightlocalizer track, with excursions over?0 feet occurring quite often even inlight wind conditions. This compareswith g feet on a two sigma basispredicted by simulation. Unfor-tunately, since the emphasis in thisprogram was on the longitudinal axis,the lateral problems were not pursuedfar enough to positively identify theirsource and resolve them.

Conclusions

The following conclusions are drawnfrom the results of these powered liftSTOP. automatic landing control lawstudies.

1. For powered lift STOL aircraftthat operate on the backside ofthe power curve, good normalacceleration control is neededfor flight path control. This

establishes requirements on bothamplitude and bandwidth. For theAugmentor Wing airplane both theengine response and the throttleservo rate limits were marginal.These limitations were partiallyovercome through the use of thedirect lift control chokes.

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2. With these automatic landingcontrol laws, the longitudinaldistance dispersion of theAugmentor Wing airplane isconsistent with STOL portrequirements ^^ defined inReference 11. these control lawsalso provide excellent sink ratecontrol.

3. The primary requirement placed onthe STOL autoland control lawdevelopment was that precise andsoft sink rata control beachijved. This is consistent withthe current practice for CTOLCategory III autoland systems.Better touchdown range control maybe possible if the allowabletouchdown sink rate is increasedthrough landing gear design or ifthe emphasis in the control lawdesign is shifted from primarilysink rate control to a combinationof sink rate and range control.

4. Good correlation was obtained inthe touchdown range and sink ratedata between flight and simulationresults through an iterativeprocess of refining mathematicalmodels and control laws per flighttest results. Under theseconditions, the fast timesimulation is effective forextrapolating the limited amountof flight data to account for lowprobability events. Additionalwork is needed to obtain similarcorrelation in the lateral axis.

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Referencesi

1. N.M. Shah, G. Gevaert, L.O. Lykken, 7. Quigley, H.C., Innis, R.C.,"The Effect of Aircraft Environment Grosswith, S., "A Flighton Category III Autoland Investigation of the STOLPerformance and Safety", AIAA 4th characterictics of an AugmentedAircraft Design; Flight Test, and Jet Flap STOL Research Aircraft",Operations Meeting, August, 1972. NASA TM X-62, 334, May 1974.

2. G. Gevaert, L.O. Lykken, N. M. 8. Scott, B.C.; Hynes, C.S.; Martin,Shah, "A Simulation Program for P.W.; and Bryder, R.B.: "ProgressCategory III Autoland Toward Development of CivilCertification" Summer Computer Airworthiness Criteria forSimulation Confernece, June 1972. Powered-Lift Aircraft" NASA

TMX-730 124 9 1976.3. Mineck, D.W., Derr, R.f., Lykken,

L.O., Hall, J.C., "Avionic Flight 9. Neuman, F., Watson, D.M.,Control System for the Lockheed Bradbury, P., "OperationalL-1011 Tristar", SAE Aerospace Description of an ExperimentalControl and Guidance Systems Digital Avionics System for STOLCommittee Meeting No. 30, September Airplanes" NASA TM X-62 9 448, 1975.1972.

10. De Hoff, R.L., Reed, W.B.,4. Anon. "Automatic Landing Systems", Trankle, T.L., Hall, Jr., W.E.,

FAA, AC 20-57A, 12 January, 1971. "Identification of Spey EngineDynamics in the Augmentor Wing Jet

5. Gevaert, G., Feinreich, B., "The STOL Research Aircraft from Flightdevelopment of Advanced Automatic Data" NASA CR-152054, October,Flare and Decrab for Powered Lift 1977.Short Haul Aircraft Using aMicrowave Landing System", NASA 11. Anon., "Planning and DesignCR- 151948, April 1977. Criteria for Metropolitan STOL

Ports", FAA Advisory Circular6. Hindson, W.S., Hardy, G.H., Innis, 150/5300-8, November 1970.

R.C., "Evaluation of Several STOLControl and Flight DirectorConcepts from Flight Tests of aPowered-Lift Aircraft Flying Steep,Curved Approaches." NASA TP 1636(To be published).

-15-

,F

' !

-SYMBOLS

aX Nozzle commanded deceleration kW Wheel command gain - deg/degrate — fps`

ay Lateral acceleration — fps 2kY Lateral position error gain -

— deg/ft

hALN Altitude where runway alignment kYl Laietal position error integralbegins — ft gain — I/see

h Vertical acceleratir>:t -- ips2 kY Lateral position rate errorgain — deg/fps

At Sink rate erro ► -- fpskya Additional lateral position rate

hFL Radar altitude at which Aare error gain at align — deg/fpssink rate control begins — ft

kbTC Throttle command gain — deg/fps

ho Altitude where runway alignmentends — ft ka^. Throttle rate command gain

— deg/sec/deghRA Radar altitude — it

lee hitch attitude gain -- degJdci119FL Radar altitude at which the

pitch Aare maneuver begins — ft k¢ t Roll integral gain — 1/sec

hTD Commanded touchdown sinkrate — fps kp Roll rate feedback gain — deg/deg/sec

kCH Choke gain — ldeg ko Yaw feedback gain — deg/deg

kh Glideslope deviation gain — fps/ft loll Yaw error integral gain — 1/see

klF Flare integrator gain —I/sec k00 Yaw to roll crossfeed gain atalign — deg/deg

klG Glideslope error integratorgain -- t/sec NH High pressure engine RPM — /o

kNCH Engine RPM to choke gain — Ic/%RPM q Pitch rate — dcg/sec

kNF RPM feedback gain — deg/ vRPM if Dynamic pressure — lb/ft2

kTF Throttle position feedback If Yaw rate — deg/secgain — deedeg

' S Laplace operator

kq pitch rate gain — deg/deg/secV

CCalibrated airspeed —fps

kr Yaw rate gain — deg/deg/sec nVC Filtered calibrated airspeed -- fps

kVl Speed integrator gain — deg/sec/fps

VREF Reference approach airspeed — fpskVN Speed gain to nozzle — deg/fps

x Longitudinal fps2acceleration —kVe Speed to pitch gain — deg/fps

i

r^

1

.1

Y Filtered lateral position error — ftAY Filtered lateral position error

rate -- ft

Y Lateral acceleration at the e.g. — fps2

TA1R Aerodynamic flight path angle — deg

AhGS Glideslope deviation -- ft

CGS Filtered glideslope deviation — ft

AkS Filtered glideslope deviationrate — fps

AYLoC Lateral position error — ft

Ali Runway heading error -- deg

A^i i Heading error when the alignmentmaneuver begins — deg

400 Heading error commanded when• the alignment maneuver ends — deg

Ti Lateral accelerometer timeconstant — see

T2 Yaw to roll crossfeed timeconstant — sec

Roll attitude — deg

Roll rate — deg/sec

Heading — deg

O RWY Runway heading — deg

dCH Choke command — %

Be Elevator command — deg

SF Flap position -{ deg

SN Nozzle angle command — deg

SN Trim table nozzle command -- deg

SNREF Reference nozzle position — deg

SK Rudder command — deg

ST Throttle position command — deg

bT Throttle rate command — deg/sec

ST Trim table throttle command — deg

6C Wheel command — deg

8 Pitch attitude — deg

CqD Touchdown pitchMOW&command — del;

8P Trim table pitch command — deg

'rCH Choke time constant — sec

Th Vertical accelerometer timeconstant — see

it

Appendix: Reading Probabilityisstri bubo Plotsis

This paper presents the flight andsimulation landing performance resultsin the form of probability curvesplotted on a graph in which a normalprobability distribution appears as astraight line. The interpretation ofthis type of probability plot ispresented We for the reader who isunfamiliar with this form of datapresentation.

In addition to determining theperformance of the system for most ofthe approaches, there is also arequirement to be sure that an unsafehard landing will be improbable. Asmatter of practice, the FAA uses the10-6 probability as the level to beassociated with the improbable event.The improbable event touchdown sinkrate from Figure 6 is 7.6 feet persecond.

Figure 5 is a curve which shows theprobability that the touchdown sinkrate exceeds the value shown on theabscissa. In Figure 6, the simulationdata probability of landing harder than2.1 feet per second is 97.7 percent.The probability of landing harder than5.5 feet per second is 2.3 percent.2.1 feet per second represents the twosigma probability of landing soft and5.5 feet per second is the two sigmaprobability of landing hard. Thedifference between these soft and hardlanding touchdown sink rates is theminus to plus two sigma sinkrate spreadand in 95.4 percent of the landings thetouchdown sink rate is between thesevalues. Thus, these limits bound themost probable performance of thecontrol system.

In this form of presentation of data, anormal probability distribution appearsas a-straight line. Non-normal datadeviate from a straight line. A verygood control system produces data whichappear as a near vertical line on theprobability graph. The poorer theperformance of the system, the more theprobability curve leans away fromvertical.

In this report, a folded probabilitycurve is used for presenting touchdowndistance data. Figure 9 is an exampleof this form of data presentation. InFiugure 9, two simulation dataprobability curves appear. Theprobability that the touchdown distanceexceeds the abscissa values is shown bythe solid line to the right of thefigure. The probability that the

touchdown range is shorter than theordinate is shown by the solid line tothe left of the figure. The curveshown on the left of the figure isobtained from the curve shown to theright by folding the top part of theP(X>XGp j p) curve vertically arouncthe mean value line. The two sigma;short landing distance can be read fromeither the P ( X>XGPIP) using the uppertwo sigma line or from the P(Wopjp)curve using the lower two sigma line.This folding of the probability curveabout the mean value permits the shortlanding improbable event value to beread opposite the long landingimprobable event value.

-Is-