2013 3rd International Conference on Instrumentation, Communications, Information Technology, and Biomedical Engineering (ICICI-BME)

Bandung, November 7-8, 2013

Design and Implementation of Hardware-In-The-

Loop-Simulation for UAV Using PID Control
Method
Sufendi1, Bambang Riyanto Trilaksono2, Syahron Hasbi Nasution3 and Eko Budi Purwanto4
1,2,3
School of Electrical Engineering and Informatics, Institut Teknologi Bandung, Bandung, Indonesia
(Tel : +62-22-202-1423; E-mail: sufendi_lie@yahoo.com)
4
Department of Applied Aerospace Technology, LAPAN, Serpong, Indonesia
(Tel : +62-21-757-90383; E-mail: ekobudi2634@binus.ac.id)

One of fixed-wing aircraft type is the fixed-wing unmanned platform to research on. The ArduPilot is one of them[13,14].
aircraft or fixed-wing Unmanned Aerial Vehicle (UAV). The UAV The tuning process of PID controller, whereby the optimum
flies without a pilot in the aircraft. All aircraft movements are values for the controller parameters are obtained, is a critical
controlled by an embedded computer or a remote control. challenge. Many studies were conducted to find the best way to
The whole complex UAV control system can be decomposed
tune PID parameters in order to get adequate performances
into several separated sections to simplify the control design
process. The three-dimensional position control was simplified to such as fast response, zero steady-state error, and minimum
one-dimensional height control and two-dimensional navigation overshoot/undershoot [3,11].
control. With educational and research purposes in autopilot control
The UAV motion composed of three force components and systems development area, a test platform is herein proposed. It
three moment components that make up the six Degrees of employs Matlab/Simulink to run the autopilot controller under
Freedom (6 DoF). In the modeling process, all motions of the test, the flight simulator X-Plane with the aircraft to be
aircraft are considered linear and have multiple inputs and commanded, a microcontroller to command model aircraft
outputs (MIMO). The controller used was a PID controller which flight control surfaces, and a servo to drive these control
is tuned using Ziegler-Nichols method. Implementation of the
surfaces[1,9,10]. In many applications relevant to the UAV, it
Hardware-in-the-Loop-Simulation (HILS) can be done after the
design of control systems for UAV is completed. The design of is necessary to convert the Global Positioning System (GPS)
HILS structure was done with the help of MATLAB software. coordinates of latitude, longitude, and altitude to a local
The controller that has been designed previously was navigation frame with coordinates : east, north, and down[5].
implemented into the Ardupilot mega hardware.
The design and implementation of PID control system with II. MODELING OF UAV
Ziegler-Nichols tuning method for the longitudinal and lateral
directional dimensions, which include angular rate control, The discussion in this section refers to the report written by
attitude control, altitude control and navigation control was able Syahron. Before conducting modeling of the UAV, perhaps
to stabilize an unstable system or to improve the system response.
general explanations about the plane should be given. A
Keyword: HILS, Ardupilot, Unmanned Aerial Vehicle, Altitude conventional fixed-wing aircraft flight control surfaces consists
control, Attitude control. of aileron, rudder, and elevator. Aileron causes the aircraft to
roll, rudder causes the aircraft to yaw, and elevator causes the
I. INTRODUCTION aircraft to pitch around the center of gravity (CG).

Unmanned Aerial Vehicles (UAVs) has been an integral part

of the military in the recent years. Besides providing numerous
advantages, one important capability of the UAVs is the ability
to carry out surveillance mission while flying autonomously,
making them powerful surveillance tools. In the process of
designing a UAV, simulation is needed to observe the attitude
and movement of UAV before it is actually implemented in a
real situation to avoid mistakes which may be fatal. UAV
Fig. 1. UAV body axis [22]
design process can be divided into various stages : form of
aircraft design stage, mathematical modeling of the aircraft, The UAV motion consists of three force components and
electronic system design, control design, flight simulation, and three moment components which form six Degree of Freedom
test flight. (6 DoF). In the modeling process, all motion of planes are
UAVs rely on autopilot to perform autonomously. Open considered linear and has multiple inputs and outputs (MIMO).
source autopilots offer a cheap and highly customizable


 


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 2013 3rd International Conference on Instrumentation, Communications, Information Technology, and Biomedical Engineering (ICICI-BME)
Bandung, November 7-8, 2013

Linearization method of equations of motion using small

perturbation theory, with consequence that the UAV equations TABEL V
Aerodynamic Characteristic and Stability Derivatives of UAV
of motion can be separated between longitudinal dan lateral
Parameter Value Unit
directional mode with loss of coupling between them.
TABLE I
Forces and Moments Acting on the Airplane Cyp 0,02888 [per rad]
Axis Direction Name Linear Small Angular Angular Cnp -0,02589 [per rad]
Velocity Displacement Velocity Clp -0,4722 [per rad/s]
OX Forward Roll U ) 5 Cnp -0,02589 [per rad/s]
OY Right Wing Pitch V T Q Cnr -0,1487 [per rad/s]
OZ Downward Yaw W \ R
Table 4 and Table 5 show that the value of Cmq, Cnr, and
Clp may be negative. This indicates that UAV is statically-
Automatic control system is an inseparable part in the stable. Here is shown the equations of UAV in the form of state
development of UAV. This system enables the aircraft to have space for both longitudinal and lateral directional modes.
stability criteria and flight control as expected. With the Equation (1) is the equation of UAV models in longitudinal
fulfillment of these criterias, the aircraft will be able to carry modes, and the equation of UAV in lateral directional modes
out its mission effectively. is shown in (2).
Test flights had also been done several times. UAV flying
conditions reviewed in this research are as follows : ‫ݑ‬ሶ ‫ݖۍ‬
‫ݔ‬௨ ‫ݔ‬ఈ ‫ݔ‬ఏ Ͳ ‫ݑ‬ ‫ ݔ‬ఋ೐
TABEL II ߙሶ ௨ ‫ݖ‬ఈ ‫ݖ‬ఏ ‫ݖ‬௤ ‫ߙ ۑې‬ ‫ݖ‬ఋ೐
൦ ሶ ൪ ൌ ‫ێێ‬ ௏ ‫ۑ‬൦ ൪ ൅ ൦
ߠ ሾߜ
Ͳ ൪ ௘ሿ (1)
ߠ
‫Ͳێ‬
UAV Flying Conditions Ͳ Ͳ ௖ ‫ۑ‬
‫ݍ‬ሶ ‫݉ۏ‬௨ ݉ఈ ݉ఏ ݉ ‫ے‬ ‫ݍ‬ ݉ ఋ೐
Parameter Value Unit ௤

‫ݕ‬ఉ ‫ݕ‬ఝ ‫ݕ‬௣ ‫ݕ‬௥

ሶ Ͳ ‫ݕ‬ఋೝ
Velocity 26 [m/sec] ‫ܫ ۍ ې ߚۍ‬௨ Ͳ ‫ܫ‬௣ ‫ܫ‬௥ ‫ې‬
ߚ
‫݌‬ሶ ‫ێ‬ ‫߮ ۑ‬ ‫ܫ‬ఋ ‫ݖ‬ఋೝ ߜ௔
‫ ۑ ێ‬ൌ ‫݊ێ‬
‫ݎ ێ‬ሶ ‫ ێ ۑ‬ఉ
Ͳ ݊௣ ݊௥ ‫ ۑ‬൦ ‫ ݌‬൪ ൅ ൦݊ ೌ
ఋೌ
൪ ൤ ൨
݊ఋೝ ߜ௥
(2)
Angle of Attack 2 [deg] ଶ௏ ‫ۑ‬
‫߮ۏ‬ሶ ‫Ͳ ۏ ے‬ Ͳ ௛
Ͳ‫ݎ ے‬ Ͳ Ͳ
Height 120 [m]
The following equation is the result of a dynamic stability
The UAV flying configuration can be seen in Tabel 3 analysis of UAV in longitudinal modes dan lateral directional
below. modes. Matrices A and B for the UAV in longitudinal mode.
TABEL III ‫ݑ‬ሶ െͲǡͳͶͺͻ Ͳǡͳʹ͵͵ െͲǡ͵͹͹Ͳ Ͳ ‫ݑ‬ Ͳ
UAV Flying Configuration ߙሶ െͲǡ͹ͷʹͷ െ͸ǡͳ͵Ͷʹ െͲǡͲͳ͵ʹ ͲǡͻͻͶͶ ߙ െͲǡ͵Ͷ͹ͳ
൦ ሶ൪ ൌ ൦
ߠ Ͳ Ͳ Ͳ ͳ
൪ ൦ߠ ൪ ൅ ൦
Ͳ
൪ ሾߜ௘ ሿ (3)
Parameter Value Unit ‫ݍ‬ሶ ͳǡͷʹͲ͵ െʹͶǡ͵Ͷ ͲǡͲʹ͸ͷ͸ െ͸ǡͲ͸͵͹ ‫ݍ‬ െͷͺǡͻʹ

MTOW 9,4 [kg] Matrices A and B for the UAV in lateral directional mode.
c.g. position NA [%mac]
ሶ െͲǡ͹͵ͳ͵ ͲǡͲͲͳͷ െͳ Ͳǡ͵͹͹ͳ ߚ Ͳ ‫ݕ‬ఋೝ
‫ې ߚۍ‬
Mac NA [m] ‫݌ ێ‬ሶ ‫ ۑ‬ൌ ൦െͶͳǡ͹ͳͷ െͳʹǡ͸ͳͳ ʹǡͶͲ͹͹ Ͳ ‫݌‬ ‫ܫ‬ ‫ݖ‬ఋೝ ߜ௔
‫ݎ ێ‬ሶ ‫ۑ‬ ͵ͳǡͺ͵ʹ െͲǡʹͲʹʹ െͳǡͳ͸ͳ͸ Ͳ
൪ ൦ ൪ ൅ ൦ ఋೌ
‫ݎ‬ ݊ఋೌ
൪൤ ൨
݊ఋೝ ߜ௥
(4)
Engine power 3,7 [hp] ‫߮ۏ‬ሶ ‫ے‬ Ͳ ͳ ͲǡͲ͵Ͷͻ Ͳ ߮ Ͳ Ͳ

Propeller diameter 20 [inch]

The modelling of actuators was conducted by using second
order-system approach. In this case, the actuator is a DC servo
Initial analysis to assess the aerodynamic characteristics and motor. The DC servo motors are considered as linear-time
stability derivatives of UAV was performed using Digital invariant systems. The transfer function for each actuator used
DATCOM software. Examples of predicted results of UAV in the UAV is :
flying configuration are given in Table 4 and Table 5. ௒ሺ௦ሻ ଷହହǡସ
ൌ మ  (5)
௑ሺ௦ሻ ௦ ାସଵǡ଺ଷହ௦ାଷହହǡସ
TABEL IV
Aerodynamic Characteristic and Stability Derivatives of UAV III. CONTROL DESIGN AND SIMULATION
Parameter Value Unit

The relatively complex UAV control system can be broken

Cxu -0,12248 [-]
down into separate layers to simplify the control design
Czu -0,62 [-] process. The decomposition of control system can be seen in
Cmu 0 [per rad] Fig 2. The position control is basically three-dimensional,
Cmq -7,976 [per rad/s] which can be decomposed into separated one-dimensional
Cmde -0,733 [per rad/s]
altitude control layer and two-dimensional trajectory tracking
control layer (in terms of geographical position, i.e. latitude
and longitude). Since both layers could be synchronized in

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Bandung, November 7-8, 2013

time, this solution does not limit the ability of the system to
track complex three-dimensional trajectory in any way [19].

Fig. 5. Response time of pitch rate control system.

Similar to the pitch rate control method, the roll rate control
also uses PID control method. Block diagram of roll rate
Fig. 2. General Block Diagram of UAV Control System [19]. control on lateral directional modes can be seen in Fig 6.

This chapter will discuss the design of the angular rate

control system, attitude control, altitude control, and navigation
control along with the simulation. In this control design, the
rudder is left in trim condition and only the elevator and aileron
are controlled.
Ardupilot mega board Aircraft

(OPh
δe
Target Altitude
h
+
-
Pengendali ketinggian
T
+
-
Pengendali sikap (sudut
pitch)
+
-
q
Pengendali kecepatan
sudut pitch
Aktuator elevator
Aircraft dynamics
longitudinal Fig. 6. Roll rate control on lateral directional modes.
T
h
(OPh (OPh
Trajectory
planning
(OP )

Bearing +
-
Pengendali navigasi +
-
Pengendali sikap (sudut
roll)
+
-
p
Pengendali kecepatan
sudut roll
δa
Aktuator aileron

Aircraft dynamics
Based on the Ziegler Nichols tuning method, the obtained
)

parameter for roll rate control is Kp= 0.1, Ki= 0.3, and Kd=
Lateral directional
(OP δr
\
0 Aktuator rudder

Sensor 0.001. The tests were performed by giving input to the system
(OPh
(OPh
Inertial Navigation System
(INS)
in the form of a constant function of 2 degrees/second. Fig. 7
(OPh
Global positioning system
(GPS)
shows the obtained response time of the roll rate control system
()T\
(p,q,r
Inerial Measurement Unit
(IMU) has a 3.81% overshoot and 6.98 seconds settling time.
Fig. 3. Block Diagram of Control System and Feedback signal that used in
each control.

III.1. Angular Rate Control Design

The control method used to control the pitch rate is PID

control method. The block diagram of the pitch rate control on
longitudinal modes can be seen in Fig 4.

Fig. 7. Response time of roll rate control system

Fig. 4. Pitch rate control on longitudinal modes.
III.2. Attitude Control System Design
Based on the Ziegler Nichols tuning method, the obtained
parameter for pitch rate control is Kp= -0.07, Ki= -0.29, and The attitude control system is designed after the angular rate
Kd= -0.002. The tests were performed by giving input to the control system has been tuned correctly. The attitude control
system in the form of a constant function of 2 degrees/second. system on longitudinal mode is also known as pitch attitude
Fig. 5 shows the obtained response time of the pitch rate hold control. The Ziegler-Nichols tuning method is used to
control system has a 3.91% overshoot and 19.3 seconds obtained control parameters for the PID controller. The values
settling-time. are as follows : Kp = 1.0 , Ki = 0.01 , dan Kd = 0.01.

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Bandung, November 7-8, 2013

III.3. Altitude Hold Control Design

The design of altitude hold control is a continuation of the

attitude control system in a longitudinal mode. In the design
process, altitude is used as the input reference and pitch angle
as the ouput signals. The pitch angle is limited at -25 degrees to
Fig. 8. Attitudes control system of pitch angle on longitudinal modes.
15 degrees. This limitation is used to avoid stall.
The reference for flying altitude is derived from the
The tests were performed by giving input to the system in waypoint planner. In this paper, UAV is considered to fly
the form of a constant function of 2 degrees/second. Fig. 9 between the ranges of trimmed condition height of 120 meters
shows the obtained response time of the system. above sea level and the cruising speed of the UAV is
considered constant at 26 m/s. Block diagram of the altitude
hold control system can be seen in Fig 12.

Fig. 12. Block diagram of altitude hold control system.

In the process of achieving the altitude obtained from the

reference waypoint , an additional block diagram is used to
linearize the changes of flying altitude with respect to the
distance, as shown in Fig. 13.

Fig. 9. Response time of pitch attitude hold control system

In the lateral directional mode, the attitude control design

used a PID controller and tuned with Ziegler-Nichols method.
Fig. 13.Block diagram of linearization of flying altitude.
The attitude control system on lateral directional mode is also
known as wing level or zero bank angle control. The control
The constraint block diagram serves as limitations to the
parameters for the attitude control in lateral directional mode
flying altitude of the UAV. To calculate the distance between
are as follows : Kp = 0.6 , Ki = 0.002 , and Kd = 0.01.
the previous waypoint to the target waypoint, the get_distance
block diagram is used. The formula to compute the distance
between waypoints is shown in (6).
‫ ݁ܿ݊ܽݐݏ݅݀݌ݓ‬ൌ  ඥሺ݊݁‫ ݐ̴݈ܽ݌ݓ̴ݐݔ‬െ ܿ‫ݐ̴݈ܽ݌ݓ̴ݐ݊݁ݎݎݑ‬ሻଶ ൅ ሺ݊݁‫ ݃݋̴݈݌ݓ̴ݐݔ‬െ ܿ‫݃݋̴݈݌ݓ̴ݐ݊݁ݎݎݑ‬ሻଶ (6)

The target_alt block diagram serves to calculate the UAV

Fig. 10. Attitudes control system of pitch angle on latral directional modes
target altitude based on its position. The formula to calculate
the altitude change based on the waypoint position is shown in
The tests were performed by giving input to the system in (7).
ሺ௪௣ௗ௜௦௧௔௡௖௘ିଷ଴ሻ‫כ‬௢௙௙௦௘௧௔௟௧
the form of a constant function of 2 degrees/second. Fig.11 ܶܽ‫ ݐ݈ܽݐ݁݃ݎ‬ൌ ݊݁‫ ݐ݈ܽ݌ݓݐݔ‬െ ቀ ቁ (7)
௪௣௧௢௧௔௟ௗ௜௦௧௔௡௖௘ିଷ଴
shows that the obtained response time of the system of the
attitude control system in lateral directional mode has 3.22% The next wp alt is the flying altitude at the destination, wp
overshoot and 15.29 seconds settling-time. distance is the distance between the target point to the current
position of the UAV. Wp total distance is the distance
between the destination waypoint to the previous waypoint,
and the alt offset is the difference between the altitude of
destination waypoint and the previous waypoint. A value of 30
is the bearing distance. All distances and altitude are in meters.

Fig. 11. Response time of wing leveler control system

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2013 3rd International Conference on Instrumentation, Communications, Information Technology, and Biomedical Engineering (ICICI-BME)
Bandung, November 7-8, 2013

Fig. 14. Change of flight altitude to the position of UAV

Ziegler-Nichols theorem is used to determine the PID

parameters for the altitude hold, based on the overall block
Fig. 16. Bearing angle on geodetic coordinate.
diagram of the control system in the longitudinal mode with
input signal of the flying altitude to be reached and output of
the current altitude. The values of Kp = 1.12 , Ki = 0.2 , and The bearing angle calculation formula is as follows :
Kd = 0.11 is obtained.
௡௘௫௧௪௣௟௔௧ି௖௨௥௥௘௡௧௟௢௖௟௔௧
ܾ݁ܽ‫ ݃݊݅ݎ‬ൌ ͻͲ ൅ ƒ”…–ƒ ቀെ ௡௘௫௧௪௣௟௢௚ି௖௨௥௥௘௡௧௟௢௖௟௔௧ቁ (8)
The test of altitude hold control on longitudinal mode is
performed by giving input of 2 meters flying altitude with The bearing angle is often called nav bearing. The
initial condition at 0 meter. Altitude hold control system have a navigation control system used bearing angle minus yaw angle
0% overshoot and 15.12 seconds settling-time The obtained obtained from feedback as the value of bearing error, as shown
response time of the system are as follows : in (9). Yaw angle can only be used as feedback in the
navigation control if the yaw angle is considered coincide with
north direction at zero yaw angle. The navigation control
system block diagram can be seen at Fig 17.
ܾ݁ܽ‫ ݎ݋ݎݎ݁݃݊݅ݎ‬ൌ ݊ܽ‫ ݃݊݅ݎܾܽ݁ݒ‬െ ܾ݁ܽ‫( ݎ݋ݎݎ݁݃݊݅ݎ‬9)

Fig. 15. Response time of altitude hold control system

III.4. Navigation Control System Design

Fig. 17. Block diagram of navigation control system
The design of navigation control system is the continuation
of the attitude control in lateral directional mode. The bearing
In the simulation process, the output of the navigation
angle is used as the input reference. The bearing value is
control system is the roll angle as a reference to the attitude
obtained from the angle between the UAV initial position and
control. The roll angle reference is limited by a function of
the target position to the north of the earth, where the value of
limit from -45 degree to 45 degree. This limit function is
the UAV position is formatted in geodetic coordinates.
called barrier function. Ziegler-Nichols method is used to
obtain the navigation control parameters on lateral directional
mode which obtains the values of Kp = 1.1 , Ki = 0.001 , and
Kd = 0.001.
The test of the navigation control in lateral directional mode
is done by gives a constant input signal of 2 degrees. The
respone time of the closed-loop system of the navigation
system can be seen in Fig 18.

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Bandung, November 7-8, 2013

PC1

(OPh APM Planner

Ardupilot mega )T\
heading

Controller Waypoint Planner

control pitch,
control roll

control pitch,
control roll

heading
)T\
(OPh
PC2

XPlane
(OPh
)T\ Aircraft dynamics,
Visualisasi
Inertial navigation sytems

MATLAB

Fig. 20. Data structure of HILS overall.

Fig. 18. Response time of the navigation control system in lateral
directional modes. The design on the MATLAB application covers about the
design of communication, design of UAV control model, and
Fig. 18 shows that the response time characteristics has UAV position estimation. To receive and transmit data
2.54% overshoot and 11.16 seconds of settling-time. between APM2.22 and MATLAB/SIMULINK, xPC target
blocksetTM is used. The data received by
IV. HARDWARE-IN-THE-LOOP-SIMULATION MATLAB/SIMULINK application is a control signal in form
of roll angle, yaw angle, and pitch angle that is forwarded by
Design of hardware-in-the-loop simulation structure is done APM2.22 application from ardupilot mega hardware. Throttle
by linking ardupilot mega hardware, as a controller, and two signal is not used as input because the UAV speed in
computers. The first computer contains APM2.22 applications considered to be constant at 26 m/s.
that is used as interface and MATLAB application that is used Afterwards, the received data is processed in the
to run the simulation state space model of UAV and as the MATLAB/SIMULINK application, resulting outputs of UAV
position estimator, while the second computer contains XPlane attitude and UAV position. The data needed by the ardupilot
application that is used as a visualization of the UAV mega hardware in the simulation is the position data of the
movement. UAV ( longitude, latitude, altitude) and the attitude of UAV (
roll angle, pitch angle, and yaw angle). The position and
attitude data are also sent to XPlane application as a
visualization of UAV, using Embedded MATLAB function
block.
Simulation Pace block from the aerospace blockset is used to
change the simulation speed to be near to real-time simulation.
The HILS overall block diagram is show in Fig 21.

Fig. 19. Communication structure of HILS overall.

Ardupilot mega hardware communicates with computer that

contains APM2.22 application using serial communication
standard, while the communication between APM2.22
application and MATLAB application using UDP
communication standard.

Fig. 21. Block diagram of HILS overall on MATLAB/SIMULINK application.

Analysis of the control system design that have been

implemented into the ardupilot mega hardware are performed
using HILS structure that has been designed before. Tests are

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Bandung, November 7-8, 2013

performed by giving a rectangular shape waypoint with

distances between waypoint is 1150 meters and varied flight
altitude. The starting point or home of the flight test is a 0
degress longitude, 0 degrees latitude, and 100 meters above sea
level. Table 6 shows the entire waypoint to be traversed by the
UAV.
TABLE VI (b)
Entire waypoint to be travered by UAV Fig. 23. (a) Waypoint of UAV and (b) flight altitude of UAV.
waypoint Latitude Longitude Altitude
(degree North) (degree East) (meter) V. RESULTS AND CONCLUSION
Home 0 0 100
Wp1 0 0.01 200
PID contol system design with Ziegler-Nichols tuning
method for longitudinal mode and lateral directional mode that
Wp2 0.01 0.01 300
covering angular rate control, attitude control, altitude control,
Wp3 0.01 0 200 and navigation control that has been designed can improve
Wp4 0 0 100 response time characteristics of UAV, where the overshoot is
less than 10% and settling time less that 20 second.
APM2.22 application is used to write a waypoint into the
ardupilot mega hardwar, as shown in Fig. 22. ACKNOWLEDGMENT
This work has been supported by a grant sponsored of National
Institute of Aeronautics and Space (LAPAN) for years 2011-
2013.
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