EMERGENCY PROPULSION-BASED AUTOLAND SYSTEM

EMERGENCY PROPULSION-BASED AUTOLAND SYSTEM

Nicolas Fezans∗ , Maxence Gamaleri∗∗DLR (German Aerospace Center) - Institute of Flight Systems

[email protected]

Keywords: Autoland, Propulsion-Controlled Aircraft, Emergency System, Monte Carlo Simulations

Abstract

In this paper, an emergency autoland system ca-pable of landing automatically with the ATTASresearch aircraft only using engine thrust vari-ations is presented. This autoland is intention-ally kept as simple as possible, but demonstratesgood performance even in the presence of uncer-tainties and external disturbances. The perfor-mance in the presence of wind shears is shownin the paper. It is based on a previously publishedpropulsion-based control law for the inner loops.Implementing this simple emergency autoland inmodern aircraft is not challenging. Aircraft al-ready in service could also be retrofitted.

1 Introduction

Jet airplanes are usually designed to be con-trolled by means of both engines (mainly act-ing on speed/energy) and control surfaces whichare deflected to generate aerodynamic forces andmoments at several places on the airplane frame.These control surfaces are usually actuated bymeans of hydraulic actuators. Quite recently,electrical actuators have also been introduced inthe most modern jet airplanes. As these sys-tems are crucial for the control of the aircraft, si-multaneous failures affecting them must be pre-vented. Therefore, aircraft control surface actu-ation systems are implemented using highly re-dundant architectures: multiple actuators, controlsignal transmission chains and power sources.

Even though aircraft are designed such thata complete failure of the primary control effec-tors is extremely improbable, this situation can

still happen and has happened several times in thepast (e.g. Japan Airlines flight 123 near Tokyo,United-Airlines flight 232 in Sioux-City or DHLin Bagdad). As shown in the three aforemen-tioned accidents it might still be possible to con-trol the aircraft by means of thrust variations. Ofcourse the maneuverability of the aircraft is thenvery restricted, but this possibility has been in-vestigated since the early 90’s when a researchprogram on propulsion controlled aircraft (PCA)took place at NASA. In this program, a pilot as-sistance system was developed and demonstratedin simulator as well as in flight tests with severalaircraft types [1–3]. More recently developmentshave been made on PCA technologies at the DLRand a propulsion-based control law was even suc-cessfully demonstrated in flight test [4, 5]. Thedeveloped system was able to assist pilots andprovide them good chances to land successfully.During the first part of our simulator studies, pi-lots did not receive explanations on the way theyshould use the system. The goal was to checkexperimentally the affordance of the developedsystem and how fast pilots were able to figureout how to use it. Results were very encourag-ing but it appeared that some of the airline pi-lots taking part to the studies misunderstood thebasic flight dynamics principles the control lawrelies on, leading sometimes to dangerous reac-tions. After short explanations, all pilots wereable to land with acceptable touchdown condi-tions in almost all following trials.

In our opinion, this good result reproducesthe results obtained by NASA in the 90’s butstill is not fully satisfying. Indeed, both airline

1

NICOLAS FEZANS, MAXENCE GAMALERI

and military pilot trainings should not be usedto train intensively for this very remote situation.Without specific training and without making pi-lot aware of the way aircraft can be controlledby means of thrust variations, the added value forthe system is clear but much lower as it seems itcould be. Apart from that, one of the results ofthe simulator tests was that the performance ofpilots seemed to be mainly limited by the men-tal workload. This was not a surprise as the air-craft reaction is very slow and pilots must be ex-tremely attentive to predict the consequences oftheir actions (even with the the assistance of thecontrol law from [5]). This suggests a completelydifferent solution: the design of an autopilot withan autoland function. This autoland permits toget results independently of pilot understandingsof the way the system is working internally. Ad-ditionally, this will permit to get a performancelevel in the presence of disturbances that a humanpilot would never obtain.

The autoland function of the autopilot will bepresented in details in section 2. In section 3, thebehaviour of the autoland is demonstrated usingMonte Carlo simulations. Finally, conclusionsand outlook are provided in section 4.

2 Propulsion-based autoland

For the design of the propulsion-based autopilot,two main options were considered:

• the design of a standalone autopilot,which would directly command the en-gines through the power lever angles,

or:

• the design of an autopilot as an outer loopfor the pre-existing propulsion-based con-trol law (see. [5]).

As it was estimated that the second option wouldease significantly the development of the autopi-lot, this option was chosen. Of course, when im-posing a structure to a controller the reachableperformance and robustness might be reduced(compared to the unstructured case). It is latershown by simulation results that the performance

level obtained is sufficient, which validates thisinitial choice.

The global structure can be represented by theblock diagram in Fig. 1 in which the way the au-topilot uses the already existing propulsion-basedcontrol law is explicitly shown. In this figure onlythe components of the autopilot that are activatedduring autoland operations are represented insidethe “autopilot” block.

AirplaneControl

lawfrom [5]

γREF PLA L

PLA R

Glideslopecontroller

Localizercontroller

Groundtrack

controller

Aut

opilo

tInertial position and speed

ΦREF

χREF

γ, Φ, p, q, r, nz, N1 L, N1 R

Fig. 1 Control structure for automatic landing

The autoland function has to intercept therunway centerline and the glidepath, to trackthem up to the ground and to land possibly witha flare. In a previous version of this autoland [6],the deviations with respect to the runway cen-terline and the glidepath were provided by theglideslope and localizer indicators. When the air-craft is quite far from the localizer and glideslopeemitters, these indicators can be interpreted as anangular measurement of the lateral and verticaldeviations. As shown in Fig. 1 and with the aimof giving more flexibility to the pilot to choosethe approach slope and later to ease the defini-tion and realisation of a flare maneuver, the clas-sical ILS measurements were replaced by the in-ertial position and speed. Internally the referencesystem used is the WGS-84 system. Of coursea relative positioning cannot be replaced by anabsolute one without some other changes: in thepresent case it must be additionally assumed thatthe position and orientation of the runway areknown with precision and that the absolute po-sitioning is also precise enough. Note that “ab-solute positioning” does not mean GPS and onlyGPS, but could be obtained by coupling inertialplatforms with GPS and barometric+radar alti-tude or any other useful source available.

2

Emergency Propulsion-based Autoland System

2.1 The underlying control law of [5]

This autopilot is designed as an outer loop usingthe control law of [5], which is already follow-ing the references on the flight path angle and onthe roll angle. In the current paper, this controllaw, how it works and how to tune it will not bere-explained in details but a short overview is pre-sented hereafter.

2.1.1 Control law requirements

In this section, the main requirements for suc-cessful approach and landing by means of a PCAsystem are discussed with focus on how desirablethey are, how difficult it will be to reach them,and which trade-off between the performance cri-teria should be made. Obviously, classical han-dling qualities criteria are not applicable for anaircraft having a propulsion-based control law.

Longitudinal Control

The longitudinal motion is mainly composedof the phugoid mode and the short-period mode.The period of the phugoid is generally between30 and 60 seconds for transport airplanes. Thefrequency of the short-period mode depends onthe aircraft and its center of gravity location, butwould typically lie between 1.5 and 3 rad/s.

Increasing the total thrust of the engines leadsto an increase of the energy rate of the aircraft. Toreally know the effect of this additional thrust onthe movement of the aircraft, the pitch equationas well as the aerodynamics and mass character-istics of the aircraft must be known. For typi-cal configurations, a simplified reasoning can beexpressed as follows: a constant additional totalthrust ∆Tt = ∑i ∆Ti > 0 leads to a positive varia-tion of the flight path angle γ and vice-versa, i.e.∆Ti > 0⇒ ∆γ(t→∞)> 0 and ∆Ti < 0⇒ ∆γ(t→∞)< 0. This makes it possible to control the tra-jectory of the aircraft in the vertical plane.

With the typical frequencies and damping ra-tios of the short-period mode and of the phugoidas well as the typical dynamics of engines, no realchallenge is expected in designing and tuning acontrol law assisting the pilots in the control of

the flight path angle. Such a control law will ba-sically consist of controlling the phugoid (accept-able response time, good damping, and no staticerror on the flight path) while avoiding unneces-sary excitation of the short-period mode.

Note that the flight path angle cannot be con-trolled independently from the speed. In order toreach the runway the angle of descent (i.e. theflight path angle) must be controlled. Once theflight path angle is determined, there is no degreeof freedom left to select the speed.

Lateral Control

The lateral dynamics of an aircraft are com-posed of:

• the Dutch roll mode exhibiting a pair ofcomplex conjugate and stable poles withvery low damping,

• the roll mode which is aperiodic and stable,

• the spiral mode which is slow and quite of-ten slightly unstable.

As for the short-period mode, the Dutch rollmode and the roll mode are generally too fastto be significantly modified by means of the en-gines, in particular in the low-thrust domain thatwill generally be required for descent and ap-proach. However, a control law based on thrustcan easily modify the spiral mode in order toease its control by a human pilot. For this, thecoupling between yaw and roll is used: the pi-lot controls only the roll motion and the con-trol law generates a yaw motion by means ofasymmetric thrust allowing to get the induced rollcorresponding to the pilot’s commands. In pre-vious studies PCA control laws were designedto follow a reference bank angle ΦREF that wasprovided by the pilot. During the current re-search activities several other possibilities havebeen investigated in the ATTAS ground simula-tor. In particular a rate-command attitude-holdand a combination of roll rate and bank anglecommands are being tested. They are both basedon the control law presented hereafter: the differ-ence is the way pilots provide the references tothe system.

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Airplane

EngR cont.

Mixingpriorities

&Protections

EngL cont.Long. cont.

Lat. cont.

Pilotinputs ΦREF

γREF

¢N1cmd

N1Rref

N1Lref PLA L

PLA R

γ,nz

Φ,p,r N1R

N1L

N1cmdPLAsat L PLAsat R

N1sat

¢N1sat

Fig. 2 Global architecture of the propulsion-based control law of [5]

Some reasonable goals for the lateral part ofthe control law are: to permit enough maneuver-ability, to reduce pilot workload by damping thelateral dynamics, and to ensure acceptable distur-bance rejection without any action of the pilot.

2.1.2 Global architecture

The global architecture of this control law is pre-sented in Fig. 2. It is based on a cascade controlstrategy with inner loops controlling the enginesthrough commands in terms of Power Lever An-gle (PLA) and outer loops controlling the longitu-dinal and lateral motions through symmetric andasymmetric thrust. The controllers in all theseloops are based on simple PI or PID structureswith feedforward terms and antiwindup. The in-ner loops have a crossfeeding in case of satura-tion (see signals PLAsatL and PLAsatR for anti-windup). A block labeled “Mixing priorities &Protections” connects the outer loops to the innerloops by allocating longitudinal (N1cmd) and lat-eral (∆N1cmd) control actions to the two engineswhile satifying the limits for each engine. Thisleads to the two references N1Lre f and N1Rre fthat are provided to the inner loops. Althoughthis does not appear very explicitly in Fig. 2 the“Mixing priorities & Protections” block also con-nects the two outer loops by means of the anti-windup feedback signals N1sat and ∆N1sat .

Further details on all the elements of this con-trol law as well as on its tuning were already pub-lished in [5] and will not be reminded in the cur-rent paper. Results of both simulator and flighttests were included and illustrate that the controllaw follow the references on the flight path angleand on the roll angle well. The autoland systempresented in this paper take the place of the pilotin the architecture of Fig. 2, as shown in Fig. 1.

2.2 Glideslope controller

The glideslope controller of the previous au-toland version was based on the geometricalprinciple presented in Fig. 3, assuming that theglideslope indicator measurements can be con-verted approximately to a deviation angle exz.Contrary to what is shown in Fig. 3, with theusual ILS system the reference point would beat the glideslope emitter, i.e. along the run-way shortly after the threshold. The positionof the reference point shown in Fig. 3 is betteras it permits to use the same approximation andto keep small deviation angles until touchdown.Small angles lead to an almost linear relation-ship between altitude error and deviation angle(exz ≈ 0⇒ tan(exz)≈ exz), which later simplifiesslightly the tracking task. For practical reasonsthe real glideslope emitter cannot be placed un-der the runway as shown in this figure, but this ispossible when defining a virtual reference pointfor the autoland system. Consequently, this refer-ence point and the touchdown point are parame-ters which can be chosen by the pilot. The config-uration shown in Fig. 3 with a reference point farbehind the touchdown point and therefore underthe runway is the one used later throughout thepaper. Note that this does not represent the usualglideslope signal (same remark applies also forthe exy later) and is only introduced here to de-fined the variable used in the description of theautoland outer loops. For gain scheduling pur-pose, the distance to the emitter or the runway isoften used: if a virtual reference point far fromthe usual emitter position is chosen, the distanceused for the gain scheduling must be correctedadequately.

The tracking of the rectilinear glidepath with

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zRW

xRW

Referencepoint

Touchdownpoint

γNOM

exzγREF

exz

γNOM

hRW

Runway

Fig. 3 Glideslope controller geometry definition

slope γNOM leading to the desired touchdownpoint (green dashed line in Fig. 3) is realised bymeans of a simple PID controller on the deviationangle exz:

γREF = γNOM +Kp(Dlon) exz . . .

+Ki(Dlon)∫

exz +Kd(Dlon) exz ,(1)

where Dlon is the distance used for the gainscheduling (usually the distance between the air-craft and the reference point). To prevent un-desirable behaviour, this PID structure is com-pleted with saturations and antiwindup (not de-tailed here but similar to the saturations and an-tiwindup shown in [5] for the inner loops). Thetime derivative exz of the deviation angle exz canbe replaced by the expression obtained by differ-entiating analytical with respect to time (no needfor numerical differentiation):

exz =x(h−hREF)− h(x− xREF)

(x− xREF)2 +(h−hREF)2 . (2)

2.3 Localizer controller

As shown in Fig. 4, the localizer controller usesthe same principle as the glideslope controller. Inmost cases, there is no need to perform lateralmaneuvers during the final part of the approach.The localizer emitter is usually placed behind theend of the runway. This position permits to pro-vide a signal of good quality during the approachand on all positions on the runway. There is noneed for selecting any other position for the refer-ence point, therefore this position was used. Thelocalizer controller works then exactly as the pre-vious version, but uses a geometrically computed

yRW

exy

xRW

ΔχREF

Centerline

Referencepointexy

RunwayDlat

Fig. 4 Localizer controller

deviation angle signal. Note that it would be pos-sible to use the exz signal previously defined forthe glideslope controller and to use the real local-izer indicator signal in the localizer controller.

The reference χREF provided by the localizercontroller to the ground track controller is di-rectly the sum of the runway direction and a cor-rection angle ∆χREF (see Fig. 4):

χREF = ΨRW +∆χREF , (3)

with ∆χREF being the output of a gain scheduledcontroller, having a PI structure:

∆χREF = Kp(Dlat,s) exy +Ki(Dlat)∫

exy , (4)

but with Kp(Dlat,s) a gain scheduled first orderfilter which is amplifying the high frequenciesmore than the low frequencies (lead-lag form).A similar effect could have been obtained using aderivative term based on exy (also analytically dif-ferentiable) combined with a first-order low-passfilter. In practice, there might be some reasons touse one or the other formulation: here the lead-lag form was simply used but this choice was notimposed by any specific constraint.

As for the glideslope controller, the schedul-ing compensate the increase of sensitivity of thedeviation angle while approaching of the refer-ence point. Only an approximated compensationwas made but it is not required to keep the closedloop gain exactly constant all along the approachtrajectory. The output of the localizer controlleris a reference on the ground track, which is pro-vided to the ground track controller whose role isto ensure that this ground track will be followedby the aircraft. The ground track controller is pre-sented in the next section.

5


2.4 Ground track controller

The ground track controller generates a referenceon the roll angle (which will be provided to thecontrol law described in [5]) which permits tomake the aircraft turn and therefore change theorientation of its ground trajectory. The lateralpart of the autoland must ensure that this trajec-tory match the desired trajectory so that the air-craft will arrive at the chosen touchdown point.

The localizer controller (see previous section)generates a ground track reference that shouldpermit to follow or rejoin the desired trajectory.The role of the ground track controller is then tomake sure that this ground track reference will befollowed.

The ground track controller consists of twomodes which are combined in a “fuzzy-controlway” for the transition between them:

• A mode used to perform small to largechanges of ground track, which only con-sists in a highly nonlinear gain on the er-ror eχ between the reference and the cur-rent ground track. This part was kept un-changed between the regular ground trackcontroller and this variation of it. In thismode the reference ΦREF reads:

ΦREF[◦] = 10 tanh

(eχ[◦]

10

). (5)

The hyperbolic tangent introduce a strongnonlinearity leading to a smooth saturationthe generated ΦREF. Extrema for ΦREF are±10◦ and are reached at eχ ≈ ±30◦. Thislimitation to 10◦ commanded roll angle isquite restrictive but prevents entering in en-gaged turns without having the control au-thority that is required to come back to hor-izontal flight. The best value for this limitcould be computed online, depending onthe current damages of the aircraft. Herefor simplicity only this 10◦ constant valuewas considered.

• A mode used for very small corrections,which after removing the integral term isbasically a proportional controller (Kp =0.25 for eχ and ΦREF in the same unit).

The transition between these two modes is madelinearly for |eχ| ∈ [0.2◦,1.5◦], which lead to anonlinear gain of 0.25 in the interval [0.0◦,0.2◦],progressively increasing to about 1 around eχ =±1.5◦. In the intervals ±[1.5◦, (almost)30◦] thecontroller behaves as a linear proportional con-troller of gain 1.

This ground track controller is a simplifiedversion of the one directly accessible to the pilotand with different gains. The pre-existing groundtrack controller could not be reused because ahigher bandwidth was required in order to use itfor centerline tracking in the presence of externaldisturbances.

3 Autoland evaluation

In this section the behaviour of the designedemergency propulsion-based autoland system isevaluated. Of course, with the same initial con-ditions, in the nominal case, and without distur-bance the system lands systematically at the sameposition, speed, and attitude. In order to eval-uate the usability of the system under more re-alistic conditions, the aircraft configuration mustbe varied (including mass, position of the centreof gravity, and possible structural damages). In-fluence of initial conditions and external distur-bances must be analysed as well. In this paper,the results shown will focus on the behaviour inthe presence of low altitude wind shears.

Although it is not shown in the paper, theinfluence of aircraft configuration as well aspossible structural damages have also been in-vestigated and the system was able to controlthe aircraft and land with all configurations andvery strong damages. For instance the air-craft landed successfully with damages generat-ing a roll torque corresponding to a fourth of theaileron roll authority at that speed. It is assumedthat aircraft configuration cannot be changed,which means that the aircraft might not be in aconfiguration permitting to satisfy the usual op-erational requirements for landing (speed cannotbe changed independently from the flight pathangle). However, reaching the runway in someof the best conditions possible would still pro-

6


vide relatively good survival chances to peopleon board. Keeping the control of the aircraft andreaching the runway permits also to prevent ad-ditional deaths on ground.

3.1 Considered wind shears

3.1.1 Wind shear direction

When controlling the aircraft only by means ofsymmetric and asymmetric thrust variations themaneuverability is far from being as good as inthe normal case. Moreover the number of degreesof freedom that a pilot can control is reduced:speed and flight path angle cannot be controlledseparately. The same applies for sideslip and roll.With an aircraft satisfying the usual handlingqualities criteria, the lateral component of windshears presents a significantly lower risk than thelongitudinal component of identical magnitude.

−14000 −11000 −8000 −5000−40

−20

0

20

40

ab

ove

th

e

glid

esl

op

e

un

de

r th

e

glid

esl

op

e

He

igh

t e

rro

r [m

]

Head wind shear

Wind shear

from right

−14000 −11000 −8000 −5000

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10

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xRW

[m]

La

tera

l err

or

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left o

f th

e

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terlin

e

rig

ht o

f th

e

ce

nte

rlin

e

Head wind shear

Wind shear

from right

0° 15° 30° 45° 60° 75° 90°

Fig. 5 Comparison between the deviations in-duced by wind shears of same magnitude (∆W =10 m/s) but different directions

Although the lateral control of the aircraft isnow significantly slower and less precise than for

an aircraft with fully functioning primary controlsystem, the performed simulations show that thedeviations induced by lateral windshears is in-deed still relatively small. This is illustrated withthe simulations shown in Fig. 5, where the air-craft reactions to wind shears of the same magni-tude but different directions are compared.

It should also be noticed that the lateral de-viations caused by the purely longitudinal windshear is higher than the lateral deviation causedby a purely lateral wind shear of same magnitude.With a perfectly symmetric aircraft, there will beno lateral deviation at all. However, no aircraftis perfectly symmetric and therefore a deviationwill be observed. Note that the effect that is ob-served here does not correspond to strong asym-metry: to fly along a straight line (i.e. χk = 0)with all control surfaces to 0◦, both a roll an-gle and a sideslip around 0.3◦ is required (im-perfect lateral trim). These relatively small val-ues are sufficient to generate the coupling shownin Fig. 5, where the head wind shear (blue line)generates the largest lateral deviations. Evenstronger couplings should be expected with dam-aged aircraft.

3.1.2 Wind shear parameterised based on alti-tude or on time?

Physically, wind shears are differences of windspeeds and directions depending on the posi-tion considered (in 3 d.o.f.). Weather conditionsevolve with time, which introduces the time as afourth dimension in the description. When con-sidering a specific landing on a specific runway ata specific time and with a specific approach path,all the wind possibilities in this 4-dimensionalspace will not be encountered: only the windalong the trajectory will affect the aircraft. Con-sequently, wind shears are often described in re-duced forms. A typical form used to describe awind shear is to define a wind speed and orienta-tion depending only on the altitude. Note that ingeneral wind shear related to microbursts cannotbe reduced to a single dimension without losingtoo much information: microburst-induced windshears are not considered in this paper.

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When describing wind shears as a function ofthe altitude, it might be observed that the posi-tive variation of winds in the direction of the air-craft speed (i.e. that causes initially a reductionof aircraft air speed) lead to stronger reactionsthan wind shears in the opposite direction. Thereason for that is well-known: when the aircraftencounters this type of wind shear, it loses liftand begins to sink quicker, which lead to an evengreater time-variation of the wind speed. Tailwind shears are thus amplified by the aircraft mo-tion they induce, whereas head wind shears areattenuated by the induced aircraft motion.

During design phases it might be useful toconsider time-based disturbances, but this ratherartificial decoupling between aircraft motion anddisturbance should be used with extreme cautionwhen doing a performance assessment. In thefollowing only altitude-based wind shears will beconsidered.

3.1.3 Wind shear distribution used

Due to the reasons given in the previous sectionsthe following wind shear distribution was chosen:

• all wind shears are in the direction of therunway,

• all wind shears are single ramps based onaltitude (see Fig. 6) with the parameters:

– ∆W : uniform on [−15 , +15] m/s,

– Hlow: uniform on [10 , 310] m,

– ∆H: uniform on [10 , 210] m.

Win

d ve

ctor

s at v

ario

us a

ltitu

des

ΔH

GroundHlow

ΔW

Fig. 6 Longitudinal single ramp wind shear

The sign and naming convention used here-after is ∆W > 0 if the variation of wind ∆

−→W goes

in the direction of the flight (i.e. ∆−→W · −→x > 0).

Therefore ∆W > 0 will be called a tail wind shearand the opposite a head wind shear. Note that thedefinition of the tail wind shear is based on thesign of the wind variation and not on the winddirection itself!

3.2 Scenario and criteria

The considered scenario for all the Monte Carlosimulation results presented hereafter is the ap-proach and landing at Hannover Airport (HAJ -EDDV) on runway 27R with the ATTAS (WFV-614) research aircraft. The simulation is alwaysinitialized with the aircraft at the coordinates52.4235◦N / 10.2806◦E and an ASL altitude of1000 metres. This initial position is located at40 km from the runway threshold, left from thecenterline and still outside the localizer receptionzone. The initial course was chosen the same asthe runway direction and the autoland will havefirst to turn right and then left to capture and holdthe centerline direction.

Though not shown in the results that are pre-sented later, this position as well as the flightdirection at initialization were varied to verifyno undesirable behaviour could be caused dur-ing switching from the regular autopilot modesto localizer capture and localizer hold. As theseswitches are working properly and the aircraftconverge first to the given trajectory, this initial-ization does not impact the results obtained whenencountering low-altitude wind shears. For thisreason, the initialization conditions are not var-ied in the analyses presented hereafter.

Four criteria are used to assess the autolandperformance

• xRW (t = T D): x-position of the touchdownpoint in the runway reference system (seered axes in Fig. 3 and 4) in x-direction,

• yRW (t = T D): y-position of the touchdownpoint in the runway reference system,

• Φ(t = T D): Roll angle at touchdown,

• h(t = T D): Sink rate at touchdown.

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The first two criteria permit to check that theaircraft landed on the runway and not too far be-hind the runway threshold to permit to stop theaircraft. Ideally the aircraft will land at the cho-sen touchdown point. Landing outside the run-way or too far behind the runway threshold isof course unacceptable regarding these criteria.Note that the limit after the threshold may be sub-ject to many discussions: in particular the safetylevel associated to any value will strongly varywith the runway length and the current brakingcapabilities of the damaged aircraft. The valueof 1250 metres behind the runway threshold waschosen.

The third criterion on the roll angle is a sim-plified version of a criteria aiming to prevent con-tact between the ground and the wings or theengines. No large roll angles were observed onthe simulations and no problem is expected here,therefore there is no real need for a more ex-act computation. This criterion is kept to de-tect rapidly any deterioration: for instance a newtuning of the autoland gains could lead to exces-sively aggressive maneuvers in ground proximity,which is a behaviour which is dangerous, must beprevented, and could remain undetected at first ifthis criterion were not constantly monitored.

The last criterion permits to evaluate howhard the landing was and to verify that the struc-tural limits of the airplane and the landing gearsare not exceeded. The landing gears of the AT-TAS are certified for a vertical impact velocityup to 700 ft/min and therefore this limit is takenfor the landing acceptance criterion.

3.3 Results

When using the autoland based on the absoluteposition, the glidepath angle can be freely chosenand can thus be flat enough to avoid the necessityof performing a flare (assuming that the terrainaround the chosen airport is flat enough). Con-sidering the current speed of the aircraft a flight-path angle value that would lead to acceptablesink rate can be computed. This value is expectedto be lower or equal to −1.5◦ with typical valuesof transport airplanes. However, this simple rea-

soning do not take into account the effect of pos-sible disturbances on the result and in particularit might be interesting to keep some margins andtherefore to choose an even more gentle slope.If this is the case, how large should these marginsbe? In order to select the slope that offers the bestchances for successful landing a detailed analysismust be performed.

The results obtained using the distributiondescribed in section 3.1.3 for two different ap-proach slopes are shown in Fig. 7. In all thesesimulations the autoland was set to follow a rec-tilinear slope until touchdown, i.e. without per-forming a flare. The upper-left plot in this fig-ure represents the proportion of accepted land-ings (using the criteria mentioned in section 3.2)for each of the simulated approach slopes. Thelower-left plot shows the respective proportionsof unaccepted landings caused by tail and headwind shears. In the upper-right plot of this fig-ure the maximum altitude at which a wind shearleading to an unaccepted landing started is rep-resented for each slope and wind shear direction(tail or head). Finally in the lower-right part ofthis figure the respective proportions between thecauses for nonacceptance are represented.

As no flare was performed, the acceptanceof landings with the slope of −2.9◦ is very low,which is logical as this slope leads to an exces-sive nominal sink rate. This acceptance increasessignificantly with the use of less steep slopes.As the nominal sink rate was already unaccept-able for the −2.9◦ slope, the results were almostidentical for head and tail wind shears: with themost gentle approach slopes differences can beobserved and tail wind shears are significantlymore dangerous. At the same time, the risk ofimpacting the ground ahead of the runway thresh-old increases drastically when using more gentleslopes. This is mainly due to the initial loss ofaltitude that tail wind shears induce. Note that inthe simulation environment “a perfectly flat air-port neighbourhood” is assumed.

If the landscape around the airport permits touse a very gentle slope (between 0.4◦and 1.2◦)and if weather conditions indicates that windshear (especially tail wind shear) encounters are

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tud

e f

or

a w

ind

sh

ea

r le

ad

ing

to

an

un

acc

ep

tab

le la

nd

ing

[m

]

tail wind shears

head wind shears

−3 −2.5 −2 −1.5 −10

50

100

25

75


Un

acc

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tab

le la

nd

ing

sta

il /

he

ad

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d s

he

ars

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utio

n [

%]

tail wind shears

head wind shears

−3 −2.5 −2 −1.5 −10

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Dis

trib

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n o

fu

na

cce

pta

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lan

din

gs

x vs

. d

h/d

t [%

]

with unacceptable xRW(touchdown)

only unacceptable xRW(touchdown)

unacceptable sink rate

Fig. 7 Results obtained with various slopes

unlikely, this should be done. Pilots have accessto the obstacles around the airports on the mapsand can thus assess the risk related to the useof such a very gentle slope. This recommenda-tion applies only to landings with the proposedautoland system: gentle slopes can lead to largevariations of the touchdown position if the trajec-tory is not tracked well enough. This is likely tohappen if the same trajectory is tracked manually.

If the weather is locally more uncertain, thefirst question is of course whether the situationis less uncertain on a different airport or runway:just as for the initial choice that the pilot wouldhave made, the runway length, the safety equip-ments and expected hazard in case of runway ex-cursion must be taken into account. If the runwayis particularly long the risk of landing before therunway due to a tail wind shear could be partlyalleviated by selecting a desired touchdown pointquite far from the runway threshold. If this wouldnot be the case, it seems at first better to make aquite hard landing at an acceptable position thana “soft” landing outside the runway or too far be-hind the threshold: most hard landings do not endup with deaths whereas quite a lot of runway ex-cursions end tragically.

In order to compare the results obtained forvarious slopes, more specific simulations mustbe looked at. To explain the differences in-duced by varying the approach slope, two MonteCarlo simulations with less parameters were per-formed: one with the 1.6◦ slope and the other

with the 2.9◦ slope. In both cases a tail windshear (∆W = 10 m/s) was always considered butits position (and thus the encounter altitude) wasvaried. The results obtained are shown in Fig. 8.In the upper part of this figure each point repre-sents the touchdown position in the runway co-ordinates: the red dot-dashed line around −3400m represents the position of the threshold (aboveis ahead of the threshold). The points that corre-spond to the same slope are connected with linesin order to show how the the touchdown positionvaries with Hlow. The lower part uses the samerepresentation for the sink rate H at touchdown:the red dot-dashed line is the certified value forthe landing gears of the ATTAS.

In this figure, the curves for x at touchdownoscillate with different periods. This is a con-sequence of the excitation of the phugoid modeby the wind shear: the resulting oscillation is su-perimposed to the “ideal path” and leads to vari-ations of the touchdown positions depending onthe altitude where it was initiated. These curvesare discontinuous because there are limit casesleading to land at one oscillation if the wind shearoccurred at a given altitude, but at the previousone if the wind shear occurred at an only slightlylower altitude. These discontinuities on the x po-sition correspond to peaks on the curves of H thatgo to 0 (or slightly lower as the limit case was notperfectly found in the simulations made). Look-ing at such curves, it is obvious that consolidatedvalues (e.g. mean, standard deviations, quantiles,

10


−2800

−4000

−3400

x(t=TD)[m

]

100 200 300 400

−1400

−700

0

Hlow [m]

H(t=TD)[ft/

min]

1.6° slope2.9° slope

Fig. 8 Variations of touchdown positions for thesame tail wind shear (10 m/s) depending on theencounter altitude.

etc.) must be taken with precaution as their de-pendence to the input distributions used is verystrong. This applies of course also to the resultsshown in Fig. 7.

In this figure the oscillations shown have dif-ferent periods. Reason for that is that the timerequired to reach the ground from the time thewind shear is encountered is approximately thequotient of the encounter altitude by the sink rate:therefore for a given encounter altitude the moregentle the slope is the more time is available torestabilise the aircraft on the ideal slope. This ex-plains why the oscillation vanishes “earlier” (i.e.at lower encounter altitudes) for the 1.6◦ sloperthan for the 2.9◦ slope. The fact more gentleslopes leads to higher risk to reach the groundbefore the runway is also clearly visible in thisfigure. The mean x-touchdown value during the“first” two oscillations with the 1.6◦ slope devi-ates clearly from the −3000 m value set for theideal touchdown point, which is not the case forthe 2.9◦ slope.

When looking at the proportion of acceptablelandings (Fig. 7), the values even for relativelyflat approaches are around 80 - 90%, which stillseems quite far from the 100%−ε that one would

like to get. It should be kept in mind that the air-craft is assumed to have lost control of its con-trol surfaces which is rather unlikely to happen.This aircraft is controlled only by means of its en-gines thanks to an emergency assistance system.In addition to this already very critical scenario,low altitude wind shears are encountered duringthe approach. Under these conditions the perfor-mance reached is already very good. Beside, theimplementation of this emergency system do notrequired specific hardware: it could run on thecurrently existing flight computers and even beintegrated in a software update of the system al-ready flying.

Taking the airspeed measurement into ac-count, the wind shear could be explicitly detectedand precompensated. This would alleviate theinitial reaction of the aircraft, but not completelycancel it. Indeed, there is no possibility to dis-tribute the aircraft energy among kinetic and po-tential energy (usually possible through the hor-izontal trim and the elevators), which would berequired to get an even more efficient alleviationof the wind shear. We only can compensate en-ergy deficits or excesses.

Though not desirable, hard landings are of-ten not fatal. Landing before the runway seemsmore critical, but if the weather is uncertain andthere was no derouting possible, the most prati-cable solution would be to choose a touchdownpoint further on the runway (here it was chosenonly 400 m behind the runway threshold) whichwould significantly reduce this risk.

4 Conclusions and outlook

An emergency propulsion-based autoland systemfor the ATTAS (VFW-614) research aircraft waspresented. It consist in a simple outer loop to acontrol law previously published by the authors[5]. The performances of the complete systemwere shown in the paper with a focus on distur-bance rejection capabilities. Various additionaldevelopments have already been made and testedin desktop simulation, which includes: curvedtrajectory tracking, flare maneuvres, and a feed-forward that is activated only for strong distur-

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bances. These systems are still being developed,but already improve the good performance of theemergency propulsion-based autoland system. Itis also planned to port this autoland (and under-lying control law) for the DLR A320 ATRA re-search aircraft and at least to test it in a simulator.

The proposed system provides very good sur-vivability chances and in most cases no hull lossshould even be expected, though the failure con-dition leading to the use of this system is cur-rently classified as “catastrophic”. A typical is-sue encountered when introducing a new emer-gency system in addition to the existing ones isrelated to the possible activation of this systemwhen not required. In the case of the emergencypropulsion-based autoland system proposed inthis paper, this issue is easy to address as it isvery easy to know with certainty whether pri-mary control systems are still working or not.Only if none of them is working the system canbe used. The readiness level of this technologyis high enough to implement it in flight controlsystems and the associated costs are very low(basically a couple of additional modes in flightcontrol softwares). To our mind, assuming thatthe pursued flight safety strategy follows the “AsLow As Reasonably Practicable” (ALARP) pre-cept: it seems there is no good reason for notintroducing this emergency system (or any suit-able alternative) in transport aircraft. Besides, thecosts assosiated with a single accident (investiga-tion, legal procedures, possible responsabilities,image loss etc.) are expected to be significantlyhigher than the development costs of this addi-tional and nonmandatory function for a completefamily of modern fly-by-wire airplanes.

References

[1] Burcham, F. W. Jr., Burken, J. J., Maine, T. A.and Fullerton, C. G., Development and flight testof an emergency flight control system using onlyengine thrust on an MD-11 transport airplane.NASA/TP-97-206217, 1997.

[2] Burcham, F. W. Jr., Maine, T. A., Fullerton, C. G.and Webb, L. D., Development and flight eval-uation of an emergency digital flight control sys-

tem using only engine thrust on an F-15 airplane.NASA/TP-3627, 1997.

[3] Bull, J., Mah, R., Hardy, G., Sullivan, B., Jones,J., Williams, D., Soukup, P., Winters, J., Pilotedsimulation tests of propulsion control as backupto loss of primary flight control for a B747-400jet transport. NASA/TM-112191, 1997.

[4] de Almeida, F., Trajectory tracking with fault-tolerant flight control system: a model predictiveapproach. PhD thesis TU Braunschweig, ISBN:978-3-8322-8546-3, 2009.

[5] Fezans, N., Simple control law structure for thecontrol of airplanes by means of their engines.in Advances in Aerospace Guidance, Naviga-tion and Control, Ed. F. Holzapfel and S. Theil,Springer, ISBN 978-3-642-19816-8, 2011.

[6] Fezans, N. and Kappenberger, C., Approachesand vista on flight control systems reconfigura-tion - motivating example: aircraft control bymeans of engine thrust. ONERA-DLR AerospaceSymposium, Toulouse, France, 2011.

[7] de Almeida, F., Waypoint navigation using con-strained infinite horizon model predictive con-trol, AIAA-2008-6462. AIAA GNC, Honolulu,HI, USA, 2008.

[8] de Almeida, F. and Leißling, D., Fault-tolerantmodel predictive control with flight test results,Journal of Guidance, Control, and Dynamics,0731-5090, Vol. 33 No. 2, 2010.

[9] Proctor, F. H., Hinton, D. A., Bowles, R. L., Awindshear hazard index, Paper: 7.7, 9th Confer-ence on Aviation, Range and Aerospace Meteo-rology, Orlando, FL, USA, 2000.

Copyright Statement

The authors confirm that they, and/or their company ororganization, hold copyright on all of the original ma-terial included in this paper. The authors also confirmthat they have obtained permission, from the copy-right holder of any third party material included in thispaper, to publish it as part of their paper. The authorsconfirm that they give permission, or have obtainedpermission from the copyright holder of this paper, forthe publication and distribution of this paper as part ofthe ICAS2012 proceedings or as individual off-printsfrom the proceedings.

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EMERGENCY PROPULSION-BASED AUTOLAND SYSTEM

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