Autolanding Controller Strategies For a Fixed Wing

UAV in Adverse Atmospheric Conditions

Ilkay Yavrucuk∗ and Volkan Kargin †

Middle East Technical University, Ankara, TURKEY

This paper focuses on the problem of autonomous landing of fixed wing UAVs in difficultatmospheric conditions, like crosswind, tail wind, turbulence. Various controller architec-tures along with landing trajectories are proposed. The experimental UAV developed atthe Middle East Technical University is used as a case study. The UAV is modeled andsimulation results are presented to demonstrate and compare the effectiveness of the de-sign architectures. Autonomous descent, flare, touch down, crab and decrab maneuversare performed using the most suitable controller design.

Nomenclature

a Acceleration, m/s2

B Input matrixg Gravity, m/s2

I Inertia of the aircraft, kgm2

K, k Controller gainL Dimensional rolling moment derivativeM Dimensional pitching moment derivativeN Dimensional yawing moment derivativen Engine revolution per secondp Roll rate, rad/sq Pitch rate, rad/sr Yaw rate, rad/su Perturbed x-body velocity, m/sv Perturbed y-body velocity, m/sw Perturbed z-body velocity, m/sX Dimensional X- force derivativeY Dimensional Y- force derivativeZ Dimensional Z- force derivativeθ Pitch angleψ Yaw angleφ Roll angleω Body angular ratesν Body velocitiesα Angle of attackβ Sideslip angleχ Aircraft velocity directionδ Control surface deflection

∗Faculty member, Department of Aerospace Engineering, yavrucuk@metu.edu.tr†Research assistant, Department of Aerospace Engineering, vkargin@ae.metu.edu.tr

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AIAA Guidance, Navigation and Control Conference and Exhibit18 - 21 August 2008, Honolulu, Hawaii

AIAA 2008-6963

Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Subscriptcom Commanded values(Input of command filters)c Commanded values(Output of command filters)cr trim values

I. Introduction

Automatic landing of an Unmanned Aerial Vehicle (UAV) is an active area of research. Since precisemaneuvering at low speeds is needed for a successful landing, it requires a combination of suitable landingstrategies, controller architectures, trajectory generators, etc. This task becomes even more challenging un-der the presence of harsh atmospheric conditions such as strong crosswinds, tailwinds, moderate to heavyturbulence, etc. Such conditions can easily challenge a UAVs landing performance and may result in catas-trophic failure when not correctly accounted for. Yet autonomous landing is desired as it provides increasedlevel of autonomy and operational consistency.

The common approach to recover a UAV is the use of a human pilot when the aircraft is in side. The pilotwould simply take over the controls of the aircraft and safely land. Another popular method is to capturethe vehicle via a safety net.1 Here the aircraft is commanded a trajectory that stirs the aircraft towards anet and the aircraft is ’captured’ in mid-air. This method is suitable for smaller UAVs and is not feasible forlarger ones. On the other hand an autonomous conventional landing of a UAV would increase performanceand could be used for larger vehicles as well.

Some recent studies investigated the problem of automatic landing of fixed wing UAVs. In Ref.2 alongitudinal controller is studied for the autolanding of a UAV. More recently H2 robust controllers weredesigned for the autonomous landing of a small fixed wing UAV.3 Similar advanced controller designs weredeveloped in Ref.4. Landing under atmospheric disturbances were also the topic of recent research: In Ref.5hardware-in-the-loop simulations are presented for the automatic landing of the Kingfisher UAV under mildcrosswinds. Similarly, in Ref.6 the automatic landing of the Heron UAV under crosswinds is investigated.

This paper makes an attempt to take a closer look on the problem, identify the key problems and proposea complete solution strategy. The design is tested under harsh atmospheric disturbances. In particular, asimulation model of the Tactical UAV designed at the Middle East Technical University is used to test andcompare controllers, trajectory generators, landing strategies. The simulations are performed under mild andstrong crosswinds, tailwinds and turbulence. While a longitudinal controller design was satisfactory for allconditions, the lateral controller design proved itself to be the most challenging. Therefore, comparisons areprovided for the controllers in the lateral channel. Autonomous crab and decrab maneuvers are performed.

The paper is organized as follows: First, the UAV and mathematical model is introduced. Then thelongitudinal and lateral autopilot designs are explained and lateral controllers are compared. Next thegeneration of the desired landing trajectory is described. Finally, simulation results for various atmosphericdisturbance scenarios are presented.

II. Mathematical Model

A tactical UAV is being developed at the Middle East Technical University as a test-bed for unmannedvehicle research. The UAV uses a 21 HP piston engine and a pusher propeller. Classic control surfaces ofthe METU TUAV are ailerons, rudder, elevator, and flaps. The engine is controlled through the throttle. Amock-up of the UAV is shown in Fig.1. Some specifications of the METU TUAV are given in Table 1.

The nonlinear simulation model of the METU TUAV has the the following major components: Nonlin-ear vehicle and control surface aerodynamics, ground effect model, propeller and engine dynamics, groundreaction and landing gear models, actuator models, atmospheric condition models, etc. The aerodynamicdatabase is created using semi-empirical formulae obtained using Refs.7 and 8. The database is updated forstall and component interaction modeling. The engine model is based on specific fuel consumption (SFC),power and RPM relationships provided by the manufacturer.9 Blade element theory (BET) is used to modelthe propeller. Dryden turbulence models and wind shear models using MIL-F-8785C10 are implemented tomodel atmospheric disturbances. Ground reaction and wheel dynamics are included to model the aircraft in

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Figure 1. METU Tactical UAV

Table 1. Specifications of the METU Tactical UAV

Wing Span: 4.3 m

Length: 1.8 m

Maximum Take-off Weight: 105 kg

Payload Weight: 25 kg

Cruise Velocity: 35-40 m/s

Maximum Range: 150 km

Maximum Endurance: 3-4 hr

Operation Altitude: 3000 m

Propulsion: 21 HP Two Cylinder Gasoline Engine

touchdown. Elevator, flaps, ailerons and rudders are modeled with second-order actuator models with rateand angular limits. The rate limits for flaps are ±100deg/s and ±200deg/s for all other control surfaces.Angular limits of elevator, rudders, ailerons and flaps are ±30deg, ±25deg, ±25deg and 0− 45deg/s, respec-tively. The engine RPM is limited between 100 and 7000rpm at standart sea level conditions but changes asa function of the air density.

The aircraft is assumed to have standard onboard sensors, which include an inertial measurement unit,GPS, compass, altimeter, airspeed measurement unit and weight on wheel sensors to detect touchdown.

III. Flight Controller Design

The proposed architecture of the controllers consist of an inner and outer loops. The inner loop controlleris designed to control the fast angular dynamics of the aircraft. Here, the control surface deflections areobtained by inverting the corresponding moment equations. The outer loop is designed to control theaircraft’s velocity and position. The proposed controller architectures use de-coupled controllers in thelateral and longitudinal channels:

A. Longitudinal Controller Design

The longitudinal controller is designed as a cascade inner and outer loop control structure. A block diagramfor the controller in the longitudinal channel is given in Fig.2.The momentum equation in the pitch channel is inverted in the inner loop to obtain the required elevatordeflection:

δe =1

Mδe + ZδeH1

[qdes − [(Mu + ZuH1)u +

Mα + ZαH1

Ucrw]− [(Mq + (Zq + Ucr)H1)q]

](1)

,where

H1 =Mα

Ucr − Zα. (2)

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Figure 2. Longitudinal controller

Here the commanded pitch angle is passed through a second order command filter. The filter parametersof the command filter are selected to match the aircraft dynamics. Then a PID controller is used to obtainthe desired pitch acceleration, θdes (Fig.3). Note that, the angular acceleration command θc is fed forwardto achieve a faster response. Similar to Ref.11 to use the inverted model of eqn.(1), the pitch acceleration,θdes, is transformed from the Euler frame to the body frame (qdes).

Figure 3. Pitch channel command filter and PID controller

In the outer loop the freestream velocity command is connected to the throttle via inverting the linearforce equation in the body x-direction. During landing it is critical to precisely hold the freestream velocityto prevent stall. The inverted force equation is:

n =1

Xn[udes − [(Xu + XTu)u +

Xα

Ucrw − g cos θcr∆θ]] (3)

A first order command filter is used in the outer loop to shape the desired velocity commands. The timeconstant in the command filter is selected as 0.15 rad/s considering the slow dynamics of the engine. Thecommand filter outputs are used in a PI controller. The desired acceleration is found from the followingequation:

Udes = Uc + Kpu(Uc − U) + Kiu

∫(Uc − U) (4)

The altitude is controlled by issuing the pitch angle command to the inner loop. A PI controller issufficient for the trajectory following during cruise, descent and climb with constant angle. However, theperformance of the controller is usually not sufficient for landing. A feed-forward term in the form of thedescent rate is used for faster convergence.12 The desired descent rate is found by:

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hcom = Ucsinγc (5)

Udes is the desired forward velocity, which is an input determined by higher order command levels and γdes

is the desired flight path angle, determined by the properties of the trajectory. Udes is taken equal to thetotal aircraft velocity. The commanded pitch angle is found from eqn 6.

θcom = Kph(hcom − h) + Kih

∫(hcom − h) + Khhcom (6)

The effect of the flaps on the aircraft is a reduction in the descent rate by producing additional lift.Contribution of flaps to the controller is limited in the descent phase. However the descent rate should beminimized in the flare phase just before touchdown. Therefore the flaps should be deflected automaticallyat flare to help the longitudinal controller.

B. Lateral Controller Design

Compared to the longitudinal controller the lateral controller design is more challenging. Several controllerarchitectures were considered for the lateral controller:

1. Lateral Controller A

The block diagram of this controller is given in Fig.4. In the inner loop the lateral momentum equations areinverted to obtain the rudder and aileron commands. The inverted equations are shown below:{

δa

δr

}=

[− (Nr+LrB1)(1−A1B1)

G1

(Lr+NrA1)((1−A1B1)G1

(Na+LaB1)(1−A1B1)G1

− (La+NaA1)((1−A1B1)G1

][

[pdes

rdes

]−

[Lβ+A1Nβ

(1−A1B1)UcrNβ+B1Lβ

(1−A1B1)Ucr

] {v}−

[Lp+A1Np

1−A1B1

Lr+A1Nr

(1−A1B1Np+B1Lp

1−A1B1

Lr+A1Nr

1−A1B1

]{p

r

}]

(7),whereG1 = −LaNr −NaA1LrB1 + NaLr + LaB1NrA1, A1 = Ixz/Ixx, B1 = Ixz/Izz

Figure 4. Lateral Controller Architecture A

Similar to the longitudinal channel, second order command filters are used for the yaw and roll channelsfollowed by PID controllers in each channel. A coordinate conversion is then performed to obtain pdes andrdes .

The more challenging part in designing the lateral channel controller appears to be in the outer loop. Inthis design the desired yaw angle for trajectory following is commanded based on the so-called cross trackerror, which is the lateral position error between the commanded trajectory and the aircraft position.

In this architecture maneuvers are accomplished using the rudder control. The primary purpose of theailerons is to keep the wings level. Therefore, the command in the roll channel is equal to zero, except a

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mixing term is added between the rudder and aileron to increase the maneuverability of the UAV in thelateral channel. So, the total aileron deflection commanded to the actuators become:

δ′a = δa + Kruδr (8)

A track guidance algorithm similar to Ref.13 is added to the controller. This guidance law requires waypoints along the landing trajectory. Here the UAV is commanded to reach the commanded way points in astraight line. However, the way points are modified through the guidance law by selecting new desired oneslocated at kxerror, where k is a user defined parameter and xerror is the previously established way point(Fig.5). If k is selected as 1, the desired location will be the original way point itself. However, it is notalways possible to achieve close trajectory following using k = 1. In that case, it is more desired to choosek < 1 and allow the aircraft to approach that way point in a straight but conditioned flight. This allows acloser following of the aircraft in presence of uncertainty, but might take longer. Note that, no parametergain scheduling or adaptation is enabled in the controller described above. Therefore k = 0.2 is selected andfound to be satisfactory in simulation and the results shown later.

Figure 5. Lateral guidance law introduced in Ref.13

In this approach the guidance law will align the ground velocity vector of the UAV with the imaginarystraight line between the UAV and the selected point (minimize ε in Fig.5). x and y are ground velocitycomponents parallel and normal to the desired path, respectively. yerror is the perpendicular distance of theUAV to the desired path. The necessary yaw rate is found by minimizing ε using the following equation:

ψcom = Ky(kxerrory − yerrorx) (9)

The term multiplied with the lateral control gain, Ky, is actually the cross product of the desired positionvector ~P and the A/C ground velocity vector ~V . Therefore the commanded yaw rate is not only a functionof ε, but also the magnitudes of ~P and ~V . Hence, the autopilot commands a higher yaw rate when the A/Cis further away from the desired position. ψcom is generated by integrating ψcom and sent into the inner loopcontroller.

2. Lateral Controller B

In this architecture, the aileron and rudder controls are used to minimize the cross track error as well as thelateral acceleration, respectively.12 The block diagram of the system is presented in Fig.6.

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Figure 6. Lateral controller architecture B

This approach makes use of the cross track error, yerror, the heading command and the aircraft velocityvector error. The aircraft velocity vector error provides a lead estimation time for the cross track error.14

Adding the aircraft heading signal provides damping in the system. As a result the reference cross trackerror becomes:

yerrorref= yerror + Uc∆t(χcom − χ) (10)

∆t is the time constant and is selected as 13s. A PI controller is used to reduce yerrorref. Note that, the

yaw rate is added as a feed forward term to achieve better tracking performance.12 Finally, the roll anglecommand is found as follows:

φcom = Kpy (yerrorref) + Kiy

∫yerrorref

+ Kψψcom (11)

Similar to lateral controller A, the angular command is sent to a second order command filter and pdes

is obtained after a linear PID controller application and coordinate transformation.The lateral acceleration command is always equal to zero. Hence no command filter is required for this

channel. Moreover a PI controller is sufficient for this channel. Finally, the corresponding inverted equationsfor the lateral dynamics are as follows:{

δr

δa

}=

[−La+NaA1

G2

Ya(1−A1B1)G2

Lr+NrA1G2

−Yr(1−A1B1)G2

][

[vdes

pdes

]−

[Yβ

UcrYp

Lβ+A1Nβ

(1−A1B1)Ucr

Lp+A1Np

1−A1B1

]{v

p

}−

[Yr − Ucr g cos θcrLr+A1Nr

1−A1B10

]{r

∆φ

}]

(12)where;

G2 = −YrLa − YrNaA1 + YaLr + NrYaA1

and ∆φ is the perturbed roll angle in the equation.

3. Lateral Controller C

The third controller architecture is designed based on Ref.15. Here an outer loop to the roll angle commandmakes use of the heading command, which is generated through another loop in the most outer layer usingthe cross track error. The yaw angle loop provides a smoother convergence and performs better to hold alevel flight condition in crosswind. The momentum equation in the roll axis is inverted in the inner mostloop. The yaw rate is controlled by the rudder and is always commanded equal to zero, a damping effect inthe yaw channel. The block diagram is presented in Fig.7.

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Figure 7. Lateral controller architecture C

The desired yaw angle is approximated using the following formula:

ψcom = ψref +1

∆tyUcyerror + Kiy

∫yerror (13)

ψref is the direction of the vector pointing to a reference point w.r.t the x-navigation axis. 1∆tyUc

is theproportional gain Ky and is determined by ∆ty and the desired velocity. ∆ty is selected as 15s. The integralcontroller is added to reduce the steady state error. It is seen that the integral gain provides benefits whenthe UAV has to follow trajectories with a certain crab angle, a case often encountered during landing incrosswinds. In Ref.15, the desired roll angle is approximated in terms of the yaw angle as follows:

φcom =Uc

∆tψg(ψcom − ψ) (14)

where g is the gravity. Note that, the value of ∆tψ determines the gain Kψ. ∆tψ is selected as 3s.

C. Comparison of Lateral Controllers

The goal is to select a controller that would be best for the aircraft to auto-land. Various simulation runsare performed to test the controllers.

Experience has shown that the lateral tracking especially under crosswinds is most critical. Therefore achallenging lateral command is given to the aircraft and the performances of the controllers are compared:A straight trajectory with length 10000m is generated between the waypoints P1(0,0) and P2(0,10000). Inthe first simulation, the UAV is initially assumed to be at PA/C(0,100). The controller feedback gain andcommand filter parameters were held fixed at their nominal values. Results are shown in Fig.8. It is observedthat controller B converged the fastest at 4000 m. The performance of the other two controllers are similarand both systems converged at about 6000m.

Next, the response of the controllers are compared for a crosswind condition of 10 m/s. Results are shownin Fig.9. A lateral error was introduced at the beginning of the simulation. The maximum cross track errorwas 10m for controller A, while for controllers B and C the maximum errors were around 40 m and 100 m,respectively. The attitude response of the aircraft is given in Fig.10. Controllers B and C kept the wingslevel after they converged to the trajectory and were able to fly with a crab angle. However, controller Bincreased the crab angle significantly since wings were not level. The cross track error signal was transmitteddirectly to the roll channel.

Finally, controller A is selected as the lateral controller for autonomous landing after comparing the threealternatives.

D. Crab and Decrab Control

Crabbing and decrabbing is a challenging maneuver during landing. During the flare maneuver the aircraftmust align itself with the local freestream velocity to achieve maximum lift, but must still fly along the

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X(m)

Y(m

)

2000 4000 6000 8000 10000-20

0

20

40

60

80

100 Lat. Controller ALat. Controller BLat. Controller CDesired Trajectory

Figure 8. Comparison of the lateral controllers following a straight trajectory(no wind)

X(m)

Y(m

)

2000 4000 6000 8000 10000-40

-20

0

20

40

60

80

100 Lat. Controller ALat. Controller BLat. Controller CDesired Trajectory

Figure 9. Comparison of the lateral controllers following a straight trajectory(10 m/s crosswind)

time(s)

phi(

deg

)

50 100 150 200 250 300

-5

0

5

10

15Lat. Controller 1Lat. Controller 2Lat. Controller 3

time(s)

psi(d

eg)

50 100 150 200 250 300-60

-55

-50

-45

-40

-35

-30

-25

-20

-15

-10

-5

0Lat. Controller 1Lat. Controller 2Lat. Controller 3

Figure 10. Comparison of yaw and roll angle controls of controllers under 10 m/s crosswind

runway, in which case the nose of the aircraft is no longer aligned with the ground velocity vector. Thecrab angle is then reduced prior to touchdown in order to align the landing gear with the runway. This iscalled decrabbing. The heading angle control is necessary at this point. As this maneuver is hard to learn

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by pilots, it is more challenging when autonomy is required.Two alternatives are investigated for the decrab maneuver: First the inner loop of lateral controller A is

used (Fig.11), i.e. the yaw angle is controlled by the rudder and the roll angle is controlled by the ailerons.The second alternative is the use of the inner loop of the controller C (Fig.12), i.e. the yaw angle is controlledby the aileron and the rudder is used to damp the yaw motion. Both approaches are tested under a 10 m/s

Figure 11. Decrab control by rudder

Figure 12. Decrab control by aileron

crosswind. At t= 250s the controller is commanded to reduce the yaw angle to 0 deg and align the headingwith the trajectory. A comparison of the performance of both approaches are shown in Fig.13. Results showthat the crab angle must be reduced by the rudder. In the first autopilot, the yaw angle is reduced from16 deg to 0 in 3 seconds with almost no overshoot. The aircraft is able to keep the wings level at the sametime. However, the second autopilot converges nearly in 10 seconds after some oscillations. The roll anglereaches 20 deg at the beginning of the maneuver. 13 deg is the limit for the aircraft to avoid hitting thewing tips to the ground. So the roll angle command should be limited during landing, which would reducethe performance of the second controller even further. Furthermore, the large deviations in the roll channelchanges the lift vector direction significantly. As a result, the UAV might not be able to create enough liftto stay in the air which would result in a sudden descent and at the best a strong impact on the landinggears.

As a result the decrab maneuver should be controlled by the rudder, as provided in the lateral controllerA.

IV. Landing Trajectory Generation

A conceptual plot of the landing trajectory for the METU TUAV is shown in Fig.14. It is assumed thatthe aircraft first descends to an altitude of 100m. The aircraft flies at this altitude for a while and alignsitself with the runway centerline and decreases its velocity. This is the point where the landing maneuverstarts. The landing maneuver can be divided into three stages: Descent, flare, touchdown and taxi.

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time(s)

psi(d

eg)

220 240 260 280 300

-10

-5

0

5decrab by ailerondecrab by rudder

time(s)

phi(

deg

)

220 240 260 280 300

0

5

10

15

20decrab by ailerondecrab by rudder

Figure 13. Response of yaw and roll angle at decrab manuever

X(m)

0

500

1000

1500

2000

2500

3000

3500

Y(m)

-0.4-0.2

00.2

0.4

Z(m

)0

20

40

60

80

100

X Y

Z

Figure 14. Conceptual landing trajectory

A. Descent Phase

The landing maneuver starts with the descent phase. The descent phase comprise the interval where theaircraft descents from 100m to the flare altitude. The glide slope and airspeed is constant during this phase.The aircraft speed and the flight path angle, γ, are the two parameters that need to be determined. A lowvelocity is preferred for a short landing distance. In addition, a high flight path angle is desired in order todecrease the approach and landing time. Unlike cruise and climb, it is not always possible to control thevelocity using only the throttle setting, since the aircraft velocity is quite sensitive to changes in pitch angleat low speeds. For instance, slowing down the aircraft might be more effective by changing the aircraft pitchangle rather than reducing the thrust. In fact, even if the engine is shut down, the aircraft might not be slowenough if a certain pitch angle is fixed. As a result the flight path angle needs to be selected not only for ashort approach time, but also for a suitable forward velocity. Note that, the longitudinal controller proposedin this paper controls the forward velocity through the throttle setting. Hence an appropriate flight pathangle selection will help to determine the aircraft speed during descent without commanding zero thrust orstalling the aircraft.

Several simulations were performed for different γ and allowable minimum forward velocity commandsbefore the controller would command a full shut down of the engine. The results are shown in Fig.15. Theallowable minimum forward velocity for flight path angles less than γ = 3.5degrees is about 25m/s. It startsincreasing after 4degrees. Considering the calculated stall speed of 23 m/s of the aircraft, the descent velocityis chosen as 28 m/s and its corresponding flight path angle at 3 degrees.

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gamma(deg)

U(m

/s)

2 2.5 3 3.5 4 4.5 5 5.5 620

30

40

50

60

Figure 15. Minimum descent velocities for different γ

B. Flare Phase

The structural loading on the landing gears should be tolerably small at touchdown. This is achieved byfurther reducing the descent rate close to the ground, called the flare phase. In Ref.16 the flare trajectoryis determined based on forward velocity, flight path angle, flare initiation altitude and flare time. Here, theforward velocity is held constant throughout the landing phase. The flare trajectory is modeled using thefollowing exponential equation:

hcom = 3.8e−t/2.5. (15)

The flare maneuver is programmed to start at about 4m altitude and is designed to last for nearly 10seconds. The flare maneuver is shown in Fig.16. The desired descent rate is given by equation 16. Thedescent rate is proportional to the altitude in this equation. It goes to zero as the altitude goes to zero.

hcom = − 12.5

hcom (16)

time(s)

Z(m

)

2 4 6 8 10

1

2

3

4

Figure 16. Flare maneuver

The taxi phase is the period after touchdown. Aerodynamic control surfaces, in particular the rudder,and the nose landing gear is used to control the heading in this phase. The throttle should be in its idleposition. The control of the taxi phase is considered to be out of the scope of this paper.

Although the trajectory and velocity control is sufficient in most of the landing maneuver the followingis worth mentioning:

In the flare phase the autopilot commands a fast pitch up maneuver to minimize altitude error whichresults in descent rate reduction. Additional lift created by flap deflection will help the A/C at this point.

Another event before touchdown is the crab angle reduction. The aircraft has to approach the runwaywith a large crab angle when under strong side winds. These angles will be larger than the angles usually

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encountered during cruise because of the low approach velocity. If the crab angle is not corrected beforetouchdown, the aircraft will not be aligned with the runway at touchdown and high lateral loads will beexperienced by the wheels. If it is reduced too early then the A/C can not keep its lateral track. Inaddition, the roll angle is also limited at touchdown to protect the wings from hitting the ground andprevent unbalanced load distribution on landing gears. After some simulations it is decided that the decrabmaneuvers should be commanded at 0.5m altitude prior to touchdown.

The automatic landing procedure is determined as shown in Fig.17. The steps are as follows:

• Approach the runway with a constant flight path angle and velocity.

• Keep lateral position error at minimum.

• Deflect flaps when flare initiates.

• Pitch up moment is created by increasing pitch angle due to reduced descent rate command.

• Reduce the crab angle and level the wings when the UAV is close to the runway.

Figure 17. Landing Procedure

V. Simulation Results

Simulation results using lateral controller A for various autonomous landing scenarios are presented next.Simulations are run for crosswinds, tail wind and turbulence scenarios. Limits for the landing maneuver forthe METU TUAV is selected as 10 m/s crosswind and 2.5 m/s tail wind.

The controllers were active from the start of the simulations until touchdown. No gain scheduling wasused within the landing maneuver. No attempt was made to control the UAV after touchdown on the ground.At touchdown, the elevator deflection is forced to its nominal value and the engine RPM is commanded toits idle position.

A. Case 1: No Wind, No Turbulence

This is the reference maneuver with no disturbances. The aircraft is commanded to land to a known runwaylocation and is assumed to be stirred to about the right location to start its landing maneuver.

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It is observed that the aircraft follows the desired trajectories very closely. The maximum error in thevertical direction is approximately 1m during the transition from level flight to descent(Fig.18) and the crosstrack error is less than 0.5m (Fig.19) at the beginning of the simulation. In Fig.20, it can be seen that flaremaneuver is successfully completed with a maximum error of 0.1m. The descent rate is reduced nearly tozero prior to touchdown. Touchdown occurs at t=115s. At touchdown instantaneous errors occur due tothe unbalance caused by the roll and heading angles. Overshoot in yaw and roll angles are quickly dampedout by ground forces in 4 seconds. The inner loop controller allows the aircraft to follow the commandsclosely(Fig.21). A reduction in forward speed due to friction forces on the landing gear can be observed inFig.19. The corresponding angle of attack, sideslip angle, body side velocity, descent rate and control surfacedeflections are provided in Figs.22, 23 and 24, respectively.

X(m)

Z(m

)

1000 1500 2000 2500 3000-20

0

20

40

60

80

100

120

A/C trajectorydesired trajectory

time(s)

Zer

ror(

m)

40 60 80 100 120-1.5

-1

-0.5

0

0.5

1

1.5

2

Figure 18. Longitudinal trajectory

X(m)

Y(m

)

1000 1500 2000 2500 3000-0.4

-0.2

0

0.2

0.4

0.6

0.8

A/C trajectorydesired trajectory

time(s)

U(m

/s)

40 60 80 100 1200

5

10

15

20

25

30

35UU desired

Figure 19. Lateral trajectory and forward velocity

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X(m)

Z(m

)

2800 2900 3000 3100 3200 3300 34000

1

2

3

4

5

A/C trajectorydesired trajectory

Figure 20. Flare maneuver

time(s)

phi(

deg

)

40 60 80 100 120-0.5

0

0.5

1

1.5

2

2.5phiphi desired

time(s)

thet

a(d

eg)

40 60 80 100 120-2

0

2

4

6

8

10thetatheta desired

time(s)

psi(d

eg)

40 60 80 100 120-1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4psipsi desired

Figure 21. Euler angles

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time(s)

alph

a(de

g)

40 60 80 100 120-2

0

2

4

6

8

10

time(s)

beta

(deg

)

40 60 80 100 120-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

Figure 22. Angle of attack and sideslip angle

time(s)

V(m

/s)

40 60 80 100 120-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

time(s)

desc

entr

ate(

m/s

)

40 60 80 100 120-3

-2.5

-2

-1.5

-1

-0.5

0

0.5

1

descent ratedesired descent rate

Figure 23. Side velocity and descend rate

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time(s)

elev

ator

defle

ctio

n(d

eg)

40 60 80 100 120-14

-12

-10

-8

-6

-4

-2

0

2

4

6

time(s)

aile

ron

defle

ctio

n(de

g)

40 60 80 100 120-16

-14

-12

-10

-8

-6

-4

-2

0

2

4

6

8

time(s)

rud

der

defle

ctio

n(de

g)

40 60 80 100 120-20

-15

-10

-5

0

5

10

15

20

time(s)

flap

def

lect

ion(

deg

)

40 60 80 100 120-2

0

2

4

6

8

10

12

time(s)

RP

M

40 60 80 100 120-1000

0

1000

2000

3000

4000

5000

Figure 24. Control surface deflections

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B. Case 2: 10 m/s Crosswind + Turbulence

A simulation scenario is performed where a crosswind of 10m/s and moderate turbulence is injected into thesimulation. The corresponding wind disturbance is given in Fig.25. The altitude and the lateral trajectory

time(s)

win

dsp

eed(

m/s

)

40 60 80 100 120 140

-4

-2

0

2

4

6

8

10

12

14 xwindywindzwind

Figure 25. Wind profiles

plots of the aircraft are given in Figs.26 and 27, respectively. The UAV approaches the flare height with nosignificant error in height or lateral position. The lateral position error is less than 2m at decrab and flaremaneuvers. The flare maneuver is initiated at t=106s. A pitch up of the aircraft is observed as expectedduring flare (Fig.28). The crab angle is seen to be around -21degrees. At a height of about 1m, ψ = 0degis commanded to reduce the crab angle (Fig.29). As a result large rudder and aileron deflections can beobserved at t=112s (Fig.32). In two seconds, the yaw angle reduces from about -20 to a value less than -1degrees and the side velocity reduces from 10 m/s to a value less than 2 m/s while the lateral controllerkeeps the roll angle at below 1deg. The decrab maneuver results in a lateral deviation of about less than 2mand loss of altitude. Touchdown occurs at t=114s. The descent rate at touchdown is about 0.5m/s tolerablefor a landing without damage(Fig.31). The corresponding angle of attack and sideslip angles are providedin Fig.30.

X(m)

Z(m

)

1500 2000 2500 3000-20

0

20

40

60

80

100

120

A/C trajectorydesired trajectory

time(s)

Zer

ror(

m)

40 60 80 100 120 140-1.5

-1

-0.5

0

0.5

1

1.5

Figure 26. Longitudinal trajectory

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X(m)

Y(m

)

1500 2000 2500 3000-10

-5

0

5

10

15

20

25

A/C trajectorydesired trajectory

time(s)

U(m

/s)

40 60 80 100 120 14022

24

26

28

30UU desired

Figure 27. Lateral trajectory and forward velocity

X(m)

Z(m

)

3000 3100 3200 3300 34000

1

2

3

4

5

A/C trajectorydesired trajectory

Figure 28. Flare maneuver

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time(s)

phi(

deg

)

40 60 80 100 120 140-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2phiphi desired

time(s)

thet

a(d

eg)

40 60 80 100 120 140-4

-2

0

2

4

6

8

10thetatheta desired

time(s)

psi(d

eg)

40 60 80 100 120 140-30

-25

-20

-15

-10

-5

0

5psipsi desired

Figure 29. Euler angles

time(s)

alph

a(de

g)

40 60 80 100 120 140-2

0

2

4

6

8

10

time(s)

beta

(deg

)

40 60 80 100 120 140-25

-20

-15

-10

-5

0

5

10

Figure 30. Angle of attack and sideslip angle

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time(s)

V(m

/s)

40 60 80 100 120 140-2

0

2

4

6

8

10

12

14

time(s)

desc

entr

ate(

m/s

)

40 60 80 100 120 140-2.5

-2

-1.5

-1

-0.5

0

0.5

1

1.5

descent ratedesired descent rate

Figure 31. Side velocity and descend rate

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time(s)

elev

ator

defle

ctio

n(d

eg)

40 60 80 100 120 140-14

-12

-10

-8

-6

-4

-2

0

2

4

6

time(s)

aile

ron

defle

ctio

n(de

g)

40 60 80 100 120 140-8

-6

-4

-2

0

2

4

time(s)

rud

der

defle

ctio

n(de

g)

40 60 80 100 120 140-30

-20

-10

0

10

20

30

time(s)

flap

def

lect

ion(

deg

)

40 60 80 100 120 140-2

0

2

4

6

8

10

12

time(s)

RP

M

40 60 80 100 120 140-1000

0

1000

2000

3000

4000

5000

6000

Figure 32. Control surface deflections

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C. Case 3: 2.5 m/s Tailwind + Turbulence

Next a landing scenario for a tail wind and moderate turbulence is performed. The wind profile used duringthe simulation is shown in Fig. 33.

time(s)

win

dsp

eed(

m/s

)

40 60 80 100

-4

-2

0

2

4

6

8

10

12

14 xwindywindzwind

Figure 33. Wind profiles

The commanded lateral and longitudinal trajectories are followed closely during all the phases of thelanding maneuver (Figs.34 and 35). Here, a higher ground velocity is commanded in order to keep airspeedconstant under the presence of tail wind. The altitude error is kept under 0.15m during the flare maneuverand the UAV touches down with a descent rate of 0.3m/s at about 95s (Figs. 36 and 39). There is nosignificant deviation at roll and yaw angles at touchdown (Fig.37). The corresponding angle of attack,sideslip angle and control surface deflections are provided in Figs.38 and 40, respectively.

X(m)

Z(m

)

1000 1500 2000 2500 3000-20

0

20

40

60

80

100

120

A/C trajectorydesired trajectory

time(s)

Zer

ror(

m)

40 60 80 100-1.5

-1

-0.5

0

0.5

1

1.5

2

2.5

Figure 34. Longitudinal trajectory

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X(m)

Y(m

)

1000 1500 2000 2500 3000-1

-0.5

0

0.5

1

1.5

2

A/C trajectorydesired trajectory

time(s)

U(m

/s)

40 60 80 10020

22

24

26

28

30

32

34

36UU desired

Figure 35. Lateral trajectory and forward velocity

X(m)

Z(m

)

2800 2900 3000 31000

1

2

3

4

5

A/C trajectorydesired trajectory

Figure 36. Flare maneuver

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time(s)

phi(

deg

)

40 60 80 100-0.2

-0.15

-0.1

-0.05

0

0.05

0.1phiphi desired

time(s)

thet

a(d

eg)

40 60 80 100-4

-2

0

2

4

6

8

10

12

14thetatheta desired

time(s)

psi(d

eg)

40 60 80 100-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8psipsi desired

Figure 37. Euler angles

time(s)

alph

a(de

g)

40 60 80 100-2

0

2

4

6

8

10

12

14

time(s)

beta

(deg

)

40 60 80 100-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

Figure 38. Angle of attack and sideslip angle

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time(s)

V(m

/s)

40 60 80 100-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

time(s)

desc

entr

ate(

m/s

)

40 60 80 100-4

-3

-2

-1

0

1

2

descent ratedesired descent rate

Figure 39. Side velocity and descend rate

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time(s)

elev

ator

defle

ctio

n(d

eg)

40 60 80 100-20

-15

-10

-5

0

5

10

time(s)

aile

ron

defle

ctio

n(de

g)

40 60 80 100-0.4

-0.2

0

0.2

0.4

0.6

0.8

time(s)

rud

der

defle

ctio

n(de

g)

40 60 80 100-0.6

-0.4

-0.2

0

0.2

0.4

0.6

time(s)

flap

def

lect

ion(

deg

)

40 60 80 100-2

0

2

4

6

8

10

12

time(s)

RP

M

40 60 80 100-1000

0

1000

2000

3000

4000

5000

6000

Figure 40. Control surface deflections

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VI. Conclusion

Controller strategies are designed, compared and simulated for a fixed wing UAV for autonomous landing.The METU TUAV is used as a case study. Controller architectures that make use of cascade inner and outerloop controllers are employed. The inner loops are used to stabilize the vehicle angular positions, while theouter loops provide the required angular position information for the corresponding position and velocitycommands. A trajectory generator provides the velocity, heading, altitude and descent rate commands for thelanding maneuver. After trying several controllers it is concluded that controlling the aircraft velocity withthe throttle, the heading with the rudder, the altitude with the elevator and using the ailerons only throughthe control mixing of the heading command, otherwise to hold the wings level was the right strategy forthe controller design. Linear controllers along with model inversion loops are used in the design. Commandfilters are employed to match the aircraft dynamics. The landing trajectory is divided into descent and flaremaneuvers. The controllers are able to follow changes in descent rate, forward speed, even quick changes fordecrab maneuvers. The controller gains and command filter parameters are all held fixed at their nominalvalues designed for slow speeds. No gain scheduling is used within the landing maneuver. Care should betaken in the coupling of the aircraft’s pitching motion -usually used to adjust the descent rate- and theaircraft’s forward speed control. The choice of an appropriate flight path angle and forward speed commandenable a safe landing without saturating the controls. The design is able to withstand various atmosphericdisturbances including crosswind, tail wind and turbulence.

References

1Lizarraga, M.I., “Autonomous Landing System for a UAV”, Naval Postgraduate School, Montarey, CA, March 2004.2Hsiao,F.B., Chan,W.L., Lai,Y.C., Tseng,L.C., Hsieh,S.Y., Tenn,H.K., “Landing Longitudinal Control System Design for

a Fixed Wing UAV” AIAA Aerospace Sciences Meeting and Exhibit, Nevada, 2007.3Rosa, P., Silvestre C., Cabecinhas, D., Cunha, R., “Autolanding Controller for a Fixed Wing Unmanned Air Vehicle”,

AIAA Guidance, Navigation and Control Conference, South Carolina, 2007.4Malaek, S.M.B., Sadati, N., Izadi, H., Pakmehr, M., “Intelligent Autolanding Controller Design using Neural Networks

and Fuzzy Logic”, 5th Asian Control Conference, 2004.5Riseborough, P., “Automatic Take-off and Landing Control for Small UAVs”, 5th Asian Control Conference, 2004.6Attar, M., Wahnon, E., Chaimovitz, D., “Advanced Flight Control Technologies for UAVs”, AIAA 2003-6537, 2003.7Hoak, D. E. and Ellison, D. E., et al. “USAF Stability and Control Datcom, Unpublished”, AF Flight Dynamics Labora-

trory, AFFDL-TR-79-3032, April 1979.8Roskam, J., “Methods for Estimating Stability and Control Derivatives of Conventional Subsonic Airplanes”, Roskam

Aviation and Engineering Corporation, 1973.9Limbach Flugmotoren website, “http://www.limflug.de/”, as accurate of 10 August, 2007.10“U.S. Military Specification MIL-F-8785C”, 5 November 1980.11Calise, A.J., Rysdyk, R.T., “Adaptive Model Inversion Flight Control for Tiltrotor Aircraft”, AIAA Guidance, Navigation

and Control Conference, August 1997.12Pashilkar, A.A., Sundararajan, N. ,Saratchandran, P., “A Fault-Tolerant Neural Aided Controller for Aircraft Auto-

landing”, Aerospace Science and Technology, Volume 10, Issue 1, January 2006.13Niculescu, M., “Lateral Track Control Law for Aerosonde UAV, AIAA 2001-0016, 2001.14Azinheira, R., de Pavia, E.C., Ramos, J.G., Bueno S.S., “Mission Path Following for an Autonomous Unmanned Airship”,

IEEE International Conference on Robotics & Automation, San Francisco, CA, April 2000.15MIT OpenCouseWare web site, “Aircraft Lateral Autopilots”, MIT URL: http://ocw.mit.edu/OcwWeb/Aeronautics-and-

Astronautics/16-333Fall-2004/, as accurate of September 1st, 2007.16McLean, D., “Automatic Flight Control Systems”, Prentice Hall, 1990.17McCormick, B. W. “Aerodynamics, Aeronautics, and Flight Mechanics”, Wiley, 1995.18Ragsdale W. A., “A Generic Landing Gear Dynamics Model For LASRS++”, AIAA 2000- 4303, 2000.19Roskam, J., “Airplane Flight Dynamics and Automatic Flight Controls”, Roskam Aviation and Engineering Corporation,

1979.

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