Design, Flight Mechanics and Flight Demonstration of a...

American Institute of Aeronautics and Astronautics

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Design, Flight Mechanics and Flight Demonstration of a Tilt-

Duct VTOL UAV

Z. Öznalbant1, M.Ş. Kavsaoğlu2, and M.Cavcar3

Anadolu University, 26470, Eskişehir, Turkey

This paper presents the design, flight mechanics and flight demonstration studies of a novel

tilt duct VTOL UAV. The aircraft, discussed in this study consists of two ducted propeller

placed on both wing tips and a ducted propeller placed between the tail booms. The aircraft

has capability of vertical take-off and landing as well as conventional take-off and landing.

Both wing tip ducts and aft duct has been designed with a capability of tilting about ninety

degree around y-axis of the aircraft. The weight estimation approach has been discussed and

initial sizes of the aircraft have been summarized. After describing the general equation of

motion, the trim condition calculations have been derived for hover, transition and cruise

flight modes. The longitudinal stability characteristics for hover, transition and cruise flights

have been analyzed via state space representation. The control strategies for all three flight

mode have been evaluated and a control algorithm has been prepared. The construction

studies of the airframe has been summarized. At the end of this study, the flight demonstration

will be completed and the comparison between the flights test and the computer simulation

results will be given.

Nomenclature

VTOL = vertical take-off and landing

CTOL = conventional take-off and landing

EoM = equations of motion

𝑚 = mass

cg = center of gravity

𝑥𝑎𝑐 = distance from cg to neutral point

𝑋, 𝑌, Z = components of resultant external force acting on aircraft

𝐿, 𝑀, 𝑁 = components of resultant external moment acting on aircraft

𝑈, 𝑉, 𝑊 = scalar components of velocity vector in body axis

𝑃, 𝑄, 𝑅 = scalar components of angular velocity vector in body axis

𝜙, 𝜃, 𝜓 = Euler angles

𝐼𝑥,𝑦,𝑧 = moments of inertia about (x, y, z)

𝐼𝑥𝑦,𝑦𝑧,𝑥𝑧 = products of inertia, (with respect to subscript)

x, y, z = body frame axes, positive x forward of AC, positive y right wing (arm), positive z downward direction

RPM = revolution per minute

PWM = pulse width modulation

MCU = micro-controller unit

IMU = inertial measurement unit

𝑊0 = gross weight

𝑊𝑒 = empty weight

𝑊𝑐𝑟𝑒𝑤 = crew weight

𝑊𝑝𝑙 = payload weight

𝑊𝑓𝑙 = fuel weight

1 Graduate Research Assistant, Faculty of Aeronautics and Astronautics, Anadolu University, 26470 Eskişehir,

Member AIAA 2 Professor, Faculty of Aeronautics and Astronautics, Anadolu University, 26470 Eskişehir, Senior Member AIAA 3 Professor, Faculty of Aeronautics and Astronautics, Anadolu University, 26470 Eskişehir, Senior Member AIAA


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𝑊𝑝𝑟𝑝 = propulsion system weight

𝑊𝑠𝑡𝑟 = structural weight

𝑉𝑐𝑟 = cruise velocity

𝑉𝑠𝑡𝑎𝑙𝑙 = stall velocity

Ckm = aerodynamic derivative coefficient of parameter k wrt parameter m

𝜇𝑚, 𝜇𝑎 = main and aft duct tilt angles

𝑇𝑚, 𝑇𝑎 = main and aft duct thrust values

𝑇𝑚𝑎𝑥𝑚 , 𝑇𝑚𝑎𝑥

𝑎 = main and aft duct max thrust values

𝛿Tn = thrust ratio of duct n

𝛿e = elevator deflection

α, β = angle of attack and side slip angle, resp.

𝑖h = horizontal stabilizer incidence angle

I. Introduction

HE rotary wing and fixed wing aircraft have their own advantages and disadvantages. The rotary wing aircraft

can take off and land vertically (VTOL) without a special runaway requirement. Additionally, since the downwash

air stream has low air temperature and velocity, rotary wing aircraft are the most suitable vehicles for search & rescue

duties. On the other hand, fixed wing aircraft have the superiority of high lift to drag ratio and high speed flying.

Although, the fixed wing aircraft have these advantages, the fixed wing aircraft need special runways for conventional

take-off and landing (CTOL). The studies in order to make together of these advantages in a fixed wing aircraft, were

started several decades ago by academicians and the industry itself1. The first aircraft which has the ability of VTOL

is a tilt rotor biplane of Henry Berliner in 19242. There were also some patent applications to the US patent office by

George Lehberger under the name of “Flying Machine”2 and by Nikola Tesla in 19283. Picirillo made a nice diagram4

of world VTOL aircraft according to their propulsion systems. Especially the behavior of the aircraft flying in

transition regime, such as velocity, nozzle exit jet velocity, nozzle angle, pitch attitude have been summarized in Ref.

1. Nowadays, because of the easy manufacturing and low cost requirements for operating5, unmanned aerial vehicles

(UAV) are generally preferred by the researchers for flight demonstration studies. There are different types of

approaches for vertical take-off and landing systems such as tilt rotor, tail sitter or tilt wing for both manned and

unmanned aerial vehicles6.

Tilt rotor VTOL UAVs is the first category for VTOL UAVs. Bell Eagle Eye, as a tilt rotor, is a good example for

industrial applications7. There is also another tilt rotor, called Smart UAV, has been designed, fabricated and tested

by KAI8. In this study, a control law has been developed and the conversion flight behavior has been investigated

experimentally. In another study done in University of Nanjing, an onboard embedded flight control systems for a tilt

rotor UAV has been developed and experimented successfully9. In this study, cycling control method has been used

during hover and transition flight phases. Israel Aerospace Industry has also a commercial tilt rotor, twin boom VTOL

UAV family called Panther and Mini Panther on the market10. TURAC VTOL UAV design study, by Özdemir et al.

is also another novel example for tilt rotor VTOL UAVs11. In this study an aircraft concept consists of two tilt rotor

placed at the leading edge of the wing and a main fan in the fuselage, has been discussed.

The tail sitter type of VTOL UAVs, especially ducted fan systems, is also another research area. Honeywell T-

Hawk™ is a good example for tail sitter ducted fan UAVs on the market12. This system has a gasoline engine and can

fly up to 40m with 8.4kg gross weight13. Liperra et al. studied the control system for a ducted fan VTOL UAV14. In

his study, a control strategy has been developed for a 9-inch diameter ducted fan aerial vehicle for not only hover

flight but also high forward speed flight. Johnson and Turbe discussed an adaptive controller design for a tail sitter

ducted VTOL UAV15. A study about the modeling the aerodynamics of ducted fan vehicles has been performed by

Ohanian, Gelhausen and Inman16.

The last type of VTOL UAV category is tilt-wing concept. The mechanical and aerodynamic design of a tandem

tilt wing UAV has been performed by Çetinsoy et al17. In this study, a quad tilt-wing concept has been discussed and

the VTOL flight tests are experienced successfully. Öner et al. studied the mathematical model of the previous tilt

wing UAV on his paper18. Sato and Muraoka have been investigated the flight controller design for a quad tilt wing

UAV19. Suziki et al., studied the attitude control of quad tilt wing UAV20. The design of gain scheduled stability and

control augmentation system for quad-tilt-wing UAV study has been performed by Tokoti et al21.

Ducted/shrouded propeller or ducted fan propulsion models is another special topic to discuss as a thrust

augmentation system. There are several studies investigating the efficiency increment of using the ducted propeller or

fan22, 23. The studies showed that preventing the losses at the wing tips and delaying the flow separation can,

T


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theoretically, increase the thrust up to 34% especially at static and low speed flights depending on the duct’s shape24,

25.

In this study, design, flight mechanics and flight demonstration studies for a novel tree ducted VTOL UAV have

been investigated theoretically and experimentally. The VTOL UAV discussed in this study consist of three ducted

propeller. Two of them are placed on the wing tips and able to rotate about ninety degrees around y axis of the aircraft.

The third ducted propeller placed between the tail booms and it is mechanically designed with the capability of rotating

ninety degree about its y axis. In each duct there are two counter rotating propeller motor. The VTOL UAV is able to

take-off and land vertically as well as conventionally. It can also translate from hover flight condition to cruise flight

condition by changing the ducts’ angle.

In the following section, the general properties of the aircraft summarized. The weight estimation approach based

on weight fractions, has been investigated and the initial sizes have been tabulated.

In the stability analysis section, the thrust ratios, elevator deflection and angle of attack values have been calculated

individually for trim condition for the hover, transition and cruise flight. The longitudinal stability analysis for all

flight phases has been performed via state space representation. In the fourth section, the control strategy has been

developed for stable flight. In the fifth section the construction studies and the flight tests have been summarized.

Conclusion and future works are given in the last section.

II. General Properties and Weight Estimation of the Aircraft

A. General Characteristics of the Unmanned Aircraft

The vertical and conventional flight with a fixed wing aircraft can be achieved by integrating the requirements of

propulsive, aerodynamic and control characteristic for both flight regimes into one aircraft properly. For VTOL

aircrafts, the poor control characteristics1, unpredictable power behavior of the engines during transition and maybe

the interference of aerodynamic forces and moments26 has to been considered for initial design phases. Several design

concepts has been evaluated such as canard, tandem or flying wing for initial studies. In this study, the Doak VZ427,

and the VTOL UAV28 designed by a research group in which one of the authors was member, are designated in

reference aircraft. Besides, according to low cost production and weight – balance properties, a conceptual sketch has

been drawn and shown in Fig. 1.

The UAV concept discussed in this study has three ducted propeller system in each of them there are two counter

rotation propeller engine. Two of ducts are placed at each wing tips and the third duct system is placed on the tail

boom. The ducted propellers can rotate ninety degrees around the y axis. The aircraft is able to take-off vertically,

perform transition to conventional flight and land vertically. The aircraft is also have capability of conventional take-

off, cruise flight and landing. Several design and control studies have been performed by the authors in the previous

periods29, 30.

Figure 1. Conceptual sketches of VTOL UAV, a) Cruise condition, b) Hover condition.


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B. Weight Estimation and Initial Sizing

The weight estimation study start with the formula given in Ref. 31.

𝑊0 = 𝑊𝑐𝑟𝑒𝑤 + 𝑊𝑝𝑙 + 𝑊𝑓𝑙 + 𝑊𝑒 (1)

In Eq. (1) the empty weight is considered as propulsion system weight and structural weight separately. Since

aircraft is an UAV and the propulsion is achieved by electrical engines, there is no crew or fuel weight. Therefore, the

formula become;

𝑊0 = 𝑊𝑝𝑙 + 𝑊𝑝𝑟𝑝 + 𝑊𝑠𝑡𝑟 (2)

Since the control and the avionic system parts can be considered for small type airframe, the 𝑊𝑝𝑙 weight in Eq. (2)

can be evaluated. The collected equipment list and their weights have been summarized in Table 1.

For the propulsion system weight estimation, several electrical motors and battery combinations have been tested

experimentally and the results are summarized in Table 2. The information given in the Table 2 is the average values

for similar type of motor – battery combinations. The last column of the Table 2 shows that the unit weight per unit

static thrust.

Table 1. Avionics and Control System Equipment Estimated Weight

Avionics and Control System Equipment Estimated Weights

# Item Quantity Unit Unit Weight, g Total, g

1 Flight Computer 1 unit 76 76,00

2 Flight Computer Power

Supplier 1 unit 173 173,00

3 Power Supplier Distributor 1 unit 145 145,00

4 Control Surface Actuators 4 unit 22 88,00

5 Receiver 1 unit 21 21,00

6 Electrical Cables 5 m 18 90,00

7 Telemetry 1 unit 18 18,00

8 GPS 1 unit 24 24,00

9 Distance Sensor 1 unit 1,2 1,25

10 IMU 1 unit 1,2 1,25

11 On/Off Switch 1 unit 8,5 8,50

12 Engine Cables 6 m 24 145,00

13 Propeller 6 unit 30 180,00

14 Duct Tilt Actuator 4 unit 64 256,00

15 Main Duct (Estimated) 2 unit 750 1.500,00

16 Aft Duct (Estimated) 1 unit 750 750,00

Total 3.477,00


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From the gathered test data Eq. (3) is derived.

𝑊prp = 0,12 ∗ 𝑇𝑆𝑡𝑎𝑡𝑖𝑐 (3)

The relation mentioned in Eq. (3) is used for propulsion system weight estimation. For the vertical take-off it is

known that the required static thrust must be greater or equal to the total weight of the aircraft. In addition to this,

there must be an excess power in order to control the aircraft during hover flight. Then, it is assumed that the required

static thrust must be greater or equal to the 1.25 of the total weight.

𝑇𝑆𝑡𝑎𝑡𝑖𝑐𝑅𝑒𝑞 = 1.25 ∗ 𝑊0 (4)

Putting the Eq. (4) into Eq. (3) yields Eq. (5).

𝑊prp = 0,12 ∗ 1,25 ∗ 𝑊0 = 0,15 ∗ 𝑊0 (5)

With the Eq. (2) and assuming𝑊e = 0,5 ∗ 𝑊0, the initial 𝑊0 estimation is calculated iteratively and the results are

tabulated in Table 3.

With initial weight estimation, the initial sizing of aircraft has been performed based on the techniques described

in Ref. 31 and the design parameters have been tabulated in Table 4.

Table 2. Propulsion system experienced thrust and weights

𝑈𝑛𝑖𝑡 𝑀𝑜𝑡𝑜𝑟

𝑆𝑡𝑎𝑡𝑖𝑐 𝑇ℎ𝑟𝑢𝑠𝑡∗, 𝑔

3 Duct (6 𝑀𝑜𝑡𝑜𝑟𝑠∗∗)𝑇𝑜𝑡𝑎𝑙

𝑇ℎ𝑟𝑢𝑠𝑡, 𝑔 𝑇𝑜𝑡𝑎𝑙 𝑊𝑒𝑖𝑔ℎ𝑡∗∗, 𝑔

𝑈𝑛𝑖𝑡 𝑊𝑒𝑖𝑔ℎ𝑡

𝑈𝑛𝑖𝑡 𝑇ℎ𝑟𝑢𝑠𝑡,

𝑔

𝑔

1.800,00 10.800,00 1.284,00 0,118888889

2.500,00 15.000,00 1.892,00 0,126133333

3.030,00 18.180,00 2.486,00 0,136743674 * The nominal voltage and the appropriate propeller suggested from dealer have been used.

** Includes the common battery weights which are easy to purchase.

Table 3. Initial 𝑾𝟎 weight estimation, iteratively.

𝑾𝟎 Guess, g 𝑾𝒆, g 𝑾𝒑𝒓𝒑, g 𝑾𝟎 Calculated, g

10.000,00 5.000,00 1.500,00 9.977,00

9.977,00 4.988,50 1.496,55 9.962,05

9.962,05 4.981,03 1.494,31 9.952,33

9.952,33 4.976,17 1.492,85 9.946,02

9.946,02 4.973,01 1.491,90 9.941,91

9.941,91 4.970,96 1.491,29 9.939,24

9.939,24 4.969,62 1.490,89 9.937,51

… … … …

9.934,33 4.967,16 1.490,15 9.934,31

9.934,31 4.967,16 1.490,15 9.934,30

9.934,30 4.967,15 1.490,15 9.934,30

9.934,30 4.967,15 1.490,14 9.934,29

9.934,29 4.967,15 1.490,14 9.934,29

9.934,29 4.967,15 1.490,14 9.934,29


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After initial weight estimation and sizing, a detailed CAD model has been prepared with the appropriate material

info. The weights of components of the aircraft have been measured from the CAD model are tabulated on Table 5.

III. Stability Analysis

A. General Equation of Motion, Aerodynamic and Thrust Models

The equations of motion (EoM) described in Ref. 32 and 33 are utilized for the trim calculations of the aircraft.

Equation (6)-(11) are the force and moment equations of the aircraft with respect to an inertial frame. X, Y, Z indicates

the forces and the L, M, N indicates the moments acting on the aircraft in the body axis system.

𝑋 − 𝑚𝑔𝑠𝑖𝑛𝜃 = 𝑚(�̇� + 𝑄𝑊 − 𝑅𝑉) (6)

𝑌 + 𝑚𝑔𝑐𝑜𝑠𝜃𝑠𝑖𝑛𝜙 = 𝑚(�̇� + 𝑅𝑈 − 𝑃𝑊) (7)

𝑍 + 𝑚𝑔𝑐𝑜𝑠𝜃𝑐𝑜𝑠𝜙 = 𝑚(�̇� + 𝑃𝑉 − 𝑄𝑈) (8)

𝐿 = 𝐼𝑥�̇� − 𝐼𝑧𝑥�̇� + 𝑄𝑅(𝐼𝑧 − 𝐼𝑦) − 𝐼𝑧𝑥𝑃𝑄 (9)

𝑀 = 𝐼𝑦�̇� − 𝑅𝑃(𝐼𝑥 − 𝐼𝑧) + 𝐼𝑧𝑥(𝑃2 − 𝑅2) (10)

𝑁 = 𝐼𝑧�̇� − 𝐼𝑥𝑧(�̇� − 𝑄𝑅) − 𝑃𝑄(𝐼𝑦 − 𝐼𝑥) (11)

In Eq. (6)-(11), forces and moments has been assumed that they consist of aerodynamic, propulsive and

gravitational components which shown in Eq. (12)

𝐹𝑜𝑟𝑐𝑒𝑠 = 𝐹𝐴 + 𝐹𝑇 + 𝐹𝐺 + 𝐹𝑜𝑡ℎ𝑒𝑟 (12) 𝑀𝑜𝑚𝑒𝑛𝑡𝑠 = 𝑀𝐴 + 𝑀𝑇 + 𝑀𝐺 + 𝐹𝑜𝑡ℎ𝑒𝑟

Table 4. Initial Sizing and Estimated Weights

Parameter Aircraft

𝑊𝑒, g 4.967

𝑊𝑃𝑟𝑝, g 1.490

𝑊𝑝𝑙, g 3.477

𝑊0, g 9.934

𝑉𝑐𝑟 , m/s 20

𝑉𝑠𝑡𝑎𝑙𝑙 , m/s 16

Parameter Wing Horizontal Stabilizer Vertical Stabilizer

S 0,71 m2 0,216 m2 0,102 m2 Aspect Ratio 8,8 3,2 1,81 c̅ (mean chord) 0,3 m 0,25 m 0,23 m

�̅�𝐴𝐶 0,25 0,327 - λ (taper ratio) 0,714 0,73 0,588 Λ𝐿𝐸 (leading edge

sweep angle) -10 der 22 der 45 der

Table 5. Measured Weights from CAD Model

Item Total, g

Fuselage Assembly 3.755,50

Wing Assembly 1.503,00

Mean Duct Assembly 2.238,00

Empennage Assembly 1.569,00

TOTAL 9.065,50


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For the aerodynamic forces and moments;

FA = [

CLqSrefSinα − CDqSrefCosαCYqSref

−CLqSrefCosα − CDqSrefSinα] (13)

MA = [

𝐶𝑙𝑞𝑆𝑟𝑒𝑓𝑏

CmqSrefc̅ CnqSrefb

] (14)

Where aerodynamic derivatives are defined34;

CL = CL0

+ CLαα + CLih

ih + CLδeδe

CD = CD0+ CDα

α

CY = Cy0+ Cyβ

β + Cyδaδa + Cyδr

δr

Cl = Cl0+ Clβ

β + Clδaδa + Clδr

δr (15)

Cm = Cm0+ Cmα

α + Cmihih + Cmδe

δe + CmqQ

Cn = Cn0+ Cnβ

β + Cnδaδa + Cnδr

δr

For the propulsive forces and moments;

𝐹𝑇 = [𝑇m𝛿T1𝐶𝑜𝑠𝜇𝑚 + 𝑇m𝛿T2𝐶𝑜𝑠𝜇𝑚 + 𝑇a𝛿T3𝐶𝑜𝑠𝜇𝑎

0−(𝑇m𝛿T1𝑆𝑖𝑛𝜇𝑚 + 𝑇m𝛿T2𝑆𝑖𝑛𝜇𝑚 + 𝑇a𝛿T3𝑆𝑖𝑛𝜇𝑎)

] (17)

𝑀𝑇 = [

(𝑇m𝛿T1𝑆𝑖𝑛𝜇𝑚 − 𝑇m𝛿T2𝑆𝑖𝑛𝜇𝑚)𝑙𝑇𝑚𝑦

(𝑇m𝛿T1𝑆𝑖𝑛𝜇𝑚 + 𝑇m𝛿T2𝑆𝑖𝑛𝜇𝑚)𝑙𝑇𝑚𝑥 − 𝑇a𝛿T3𝑆𝑖𝑛𝜇𝑎𝑙𝑇𝑎𝑥

(𝑇m𝛿T1𝐶𝑜𝑠𝜇𝑚 − 𝑇m𝛿T2𝐶𝑜𝑠𝜇𝑚)𝑙𝑇𝑚𝑦

] (18)

Where;

𝑇m = 𝑇𝑚𝑎𝑥

𝑚 −𝜕𝑇

𝜕𝑢𝑈

𝑇a = 𝑇𝑚𝑎𝑥𝑎 −

𝜕𝑇

𝜕𝑢𝑈

(19)

For the gravitational forces;

𝐹𝐺𝑥 = [

−𝑚𝑔𝑠𝑖𝑛𝜃𝑚𝑔𝑐𝑜𝑠𝜃𝑠𝑖𝑛𝜙𝑚𝑔𝑐𝑜𝑠𝜃𝑐𝑜𝑠𝜙

] (20)

If the aerodynamic, propulsive and gravitational forces and moments are put in to Eq. (6) - (11), the general

nonlinear equation of motion which are given in Eq. (21)-(26) are obtained.

𝑋: (𝐶𝐿0+ 𝐶𝐿𝛼

𝛼 + 𝐶𝐿𝑖ℎ𝑖ℎ + 𝐶𝐿𝛿𝑒

𝛿𝑒)𝑞𝑆𝑟𝑒𝑓𝑆𝑖𝑛𝛼 − (𝐶𝐷0+ 𝐶𝐷𝛼

𝛼)𝑞𝑆𝑟𝑒𝑓𝐶𝑜𝑠𝛼 + 𝑇m𝛿T1𝐶𝑜𝑠𝜇𝑚 + 𝑇m𝛿T2𝐶𝑜𝑠𝜇𝑚 +

𝑇a𝛿T3𝐶𝑜𝑠𝜇𝑎 − 𝑚𝑔𝑠𝑖𝑛𝜃 = 0 (21)

𝑌: (𝐶𝑦0+ 𝐶𝑦𝛽

𝛽 + 𝐶𝑦𝛿𝑎𝛿𝑎 + 𝐶𝑦𝛿𝑟

𝛿𝑟) 𝑞𝑆𝑟𝑒𝑓 + 𝑚𝑔𝑐𝑜𝑠𝜃𝑠𝑖𝑛𝜙 = 0 (22)

𝑍: −(𝐶𝐿0+ 𝐶𝐿𝛼

𝛼 + 𝐶𝐿𝑖ℎ𝑖ℎ + 𝐶𝐿𝛿𝑒

𝛿𝑒)𝑞𝑆𝑟𝑒𝑓𝐶𝑜𝑠𝛼 − (𝐶𝐷0+ 𝐶𝐷𝛼

𝛼)𝑞𝑆𝑟𝑒𝑓𝑆𝑖𝑛𝛼 − (𝑇m𝛿T1𝑆𝑖𝑛𝜇𝑚 + 𝑇m𝛿T2𝑆𝑖𝑛𝜇𝑚 +

𝑇a𝛿T3𝑆𝑖𝑛𝜇𝑎) + 𝑚𝑔𝑐𝑜𝑠𝜃 = 0 (23)

𝐿: (𝐶𝑙0+ 𝐶𝑙𝛽

𝛽 + 𝐶𝑙𝛿𝑎𝛿𝑎 + 𝐶𝑙𝛿𝑟

𝛿𝑟) 𝑞𝑆𝑟𝑒𝑓𝑏 + (𝑇m𝛿T1𝑆𝑖𝑛𝜇𝑚 − 𝑇m𝛿T2𝑆𝑖𝑛𝜇𝑚)𝑙𝑇𝑚𝑦 = 0 (44)

𝑀: (𝐶𝑚0+ 𝐶𝑚𝛼

𝛼 + 𝐶𝑚𝑖ℎ𝑖ℎ + 𝐶𝑚𝛿𝑒

𝛿𝑒 + 𝐶𝑚𝑞𝑄) 𝑞𝑆𝑟𝑒𝑓𝑐̅ + (𝑇m𝛿T1𝑆𝑖𝑛𝜇𝑚 + 𝑇m𝛿T2𝑆𝑖𝑛𝜇𝑚)𝑙𝑇𝑚𝑥 −

(𝑇m𝛿T1𝐶𝑜𝑠𝜇𝑚 + 𝑇m𝛿T2𝐶𝑜𝑠𝜇𝑚)𝑙𝑇𝑚𝑧 − 𝑇a𝛿T3𝑆𝑖𝑛𝜇𝑎𝑙𝑇𝑎𝑥 − 𝑇a𝛿T3𝐶𝑜𝑠𝜇𝑎𝑙𝑇𝑎𝑧 = 0 (25)

𝑁: (𝐶𝑛0+ 𝐶𝑛𝛽

𝛽 + 𝐶𝑛𝛿𝑎𝛿𝑎 + 𝐶𝑛𝛿𝑟

𝛿𝑟) 𝑞𝑆𝑟𝑒𝑓𝑏 + (𝑇a𝛿T1𝐶𝑜𝑠𝜇𝑚 − 𝑇a𝛿T2𝐶𝑜𝑠𝜇𝑚)𝑙𝑇𝑚𝑦 = 0 (26)


8

B. Trim Condition

For a steady level flight the following conditions must be sustained35, 36.

�̇� = �̇� = �̇� = �̇� = �̇� = �̇� = 𝜙 = 0

𝑃 = 𝑄 = 𝑅 = 0

With these assumptions, Eq. (6)-(11) become;

FAx + FTx + FGx = 0 (27)

FAy + FTy + FGx = 0 (28)

FAz + FTz = 0 (29)

LA + LT = 0 (30)

MA + MT = 0 (31)

NA + NT = 0 (32)

The propulsive forces acting on the aircraft, and the moment arms have been shown in the Figure 2. Figure 2 shows

the aircraft in vertical flight configuration.

For vertical flight condition trim calculations, it is assumed that;

𝜇m = 𝜇1 = 𝜇2 = 90 𝑑𝑒𝑔; 𝜇a = 𝜇3 = 90 𝑑𝑒𝑔, 𝜃 = 𝛼 = 3,45 𝑑𝑒𝑔, V∞ = 0 =>

𝛿T1 = 𝛿T2 = 𝛿Tm

From Eq. (21)-(26), 𝛿T1, 𝛿T2 and 𝛿T3 calculated iteratively with Gauss Seidel numerical method35. The results are

tabulated on Table 6.

𝛿Tmx =𝑚𝑔𝑠𝑖𝑛𝜃+𝑇a𝛿T3𝐶𝑜𝑠𝜇𝑎

2𝑇m (33)

𝛿Tmz =𝑇a𝛿T3𝑆𝑖𝑛𝜇𝑎−𝑚𝑔𝑐𝑜𝑠𝜃

2𝑇m (34)

𝛿T3 =2𝑇m𝛿Tmz𝑙𝑇𝑚𝑥

𝑇a𝑙𝑇𝑎𝑥 (35)

Where;

𝛿Tm = √𝛿T12 + 𝛿T2

2

Figure 2. Propulsive Forces and Moment Arms


9

𝜇𝑚 = 𝑡𝑎𝑛−1 (𝑆𝑖𝑛𝜇𝑚

𝐶𝑜𝑠𝜇𝑚

)

For steady level flight condition, it is assumed that;

𝜇m = 𝜇1 = 𝜇2 = 0; 𝜇a = 𝜇3 = 0 𝑑𝑒𝑟, γ = 0;

𝑙𝑇𝑚𝑧 = 𝑙𝑇𝑎𝑧 = 0;

𝜙 = 0 => 𝛿T1 = 𝛿T2 = 𝛿Tm;

𝛿T3 = 0;

From Eq. (21)-(26), 𝛿T1, 𝛿T2, 𝛿e and 𝛼 calculated iteratively with Gauss Seidel37 numerical method. The results

calculated for different flight conditions are tabulated on Table 7.

𝛿Tm = (−(𝐶𝐿0+𝐶𝐿𝛼𝛼+𝐶𝐿𝑖ℎ

𝑖ℎ+𝐶𝐿𝛿𝑒𝛿𝑒)𝑞𝑆𝑟𝑒𝑓𝑆𝑖𝑛𝛼+(𝐶𝐷0+𝐶𝐷𝛼𝛼)𝑞𝑆𝑟𝑒𝑓𝐶𝑜𝑠𝛼+𝑚𝑔𝑠𝑖𝑛𝜃)

2𝑇𝑚 (36)

𝛼 =(𝑚𝑔𝑐𝑜𝑠𝜃−(𝐶𝐿0+𝐶𝐿𝑖ℎ

𝑖ℎ+𝐶𝐿𝛿𝑒𝛿𝑒)𝑞𝑆𝑟𝑒𝑓𝐶𝑜𝑠𝛼−𝐶𝐷0𝑞𝑆𝑟𝑒𝑓𝑆𝑖𝑛𝛼)

(𝐶𝐿𝛼𝐶𝑜𝑠𝛼+𝐶𝐷𝛼𝑆𝑖𝑛𝛼)𝑞𝑆𝑟𝑒𝑓 (37)

𝛿𝑒 =−(𝐶𝑚0+𝐶𝑚𝛼𝛼+𝐶𝑚𝑖ℎ

𝑖ℎ+𝐶𝑚𝑞𝑄)

𝐶𝑚𝛿𝑒

(38)

For the transition flight mode, the several constrains have been defined. The first constrain is that the aircraft must

sustain its horizontal attitude and the pitch angle must be kept equal to the steady level flight pitch angle. The second

constrain is that pitch attitude is controlled by the aft engine and elevator deflection together with a linear allocation

factor. The allocation factor expressed as

𝜂 = {

1𝑉−𝑉1

𝑉2−𝑉1

0

𝑖𝑓 0𝑚/𝑠 ≤ 𝑉 < 10𝑚/𝑠𝑖𝑓 10𝑚/𝑠 ≤ 𝑉 ≤ 15𝑚/𝑠

𝑖𝑓 15𝑚/𝑠 < 𝑉 ≤ 15𝑚/𝑠 (39)

Where;

𝑉1 = 10 𝑚/𝑠 𝑎𝑛𝑑 𝑉2 = 15 𝑚/𝑠

𝛿T1, 𝛿T2, 𝛿T3, and 𝛿e calculated iteratively with Gauss Seidel37 numerical method for transition mode. Starting

from 0.1m/s to 20m/s, each trim condition has been calculated for 0.1m/s velocity increments. For the trim

calculations, it is assumed that the X components of the thrust vector must be equal to components of the aerodynamic

and the gravitational forces in the opposite direction along the X axis. In addition, along the Z axis, it is assumed that

the sum of the lift and z components of the thrust vector must be equal to gravitational forces components in the

opposite Z direction. Meanwhile, the zero pitching moment is assumed to be sustained by the aft engine and elevator

deflection. After calculating the Eq.s (40)-(46), the trim conditions have been gathered and tabulated in Table 8.

For transition flight trim calculations, the pitching moment equation divided into two part in order to calculate the

𝛿T3, and 𝛿e properly. From the hover position to 10m/s flight velocity, it is assumed that the pitch attitude is controlled

only by the aft duct; and from 15m/s to 20m/s, the pitch attitude is controlled only by the elevator deflection. Between

Table 6. Hover Flight Trim Values

𝛿T1 𝛿T2 𝛿T3 𝛿e

0,7745 0,7745 0,5963 0,0

Table 7. Cruise Flight Trim Values for Various Flight Conditions

Altitude, m V∞, 𝑚/𝑠 𝛼, 𝑑𝑒𝑔 𝛿T1 𝛿T2 𝛿T3 𝛿e

50 16 7,65 0,09 0,09 0,0 -13,13

50 18 5,29 0,09 0,09 0,0 -11,81

50 20 3,49 0,102 0,102 0,0 -10,86


10

the 10m/s to 15m/s, the pitch attitude will be controlled both aft engine and elevator deflection with an allocation

factor which was described in Eq. (39).

𝛿Tmx = 𝛿Tm𝐶𝑜𝑠𝜇𝑚 = −

(

(𝐶𝐿0+𝐶𝐿𝛼𝛼+𝐶𝐿𝑖ℎ𝑖ℎ+𝐶𝐿𝛿𝑒

𝛿𝑒)𝑞𝑆𝑟𝑒𝑓𝑆𝑖𝑛𝛼−

(𝐶𝐷0+𝐶𝐷𝛼𝛼)𝑞𝑆𝑟𝑒𝑓𝐶𝑜𝑠𝛼+

𝑇a𝛿T3𝐶𝑜𝑠𝜇𝑎−𝑚𝑔𝑠𝑖𝑛𝜃

)

2𝑇m (40)

𝛿Tmz = 𝛿Tm𝑆𝑖𝑛𝜇𝑚 =

(

−(𝐶𝐿0+𝐶𝐿𝛼𝛼+𝐶𝐿𝑖ℎ𝑖ℎ+𝐶𝐿𝛿𝑒

𝛿𝑒)𝑞𝑆𝑟𝑒𝑓𝐶𝑜𝑠𝛼−

(𝐶𝐷0+𝐶𝐷𝛼𝛼)𝑞𝑆𝑟𝑒𝑓𝑆𝑖𝑛𝛼−

𝑇a𝛿T3𝑆𝑖𝑛𝜇𝑎+𝑚𝑔𝑐𝑜𝑠𝜃

)

2𝑇m (41)

𝑅𝐻𝑆𝑀 = (𝐶𝑚0+ 𝐶𝑚𝛼

𝛼 + 𝐶𝑚𝑖ℎ𝑖ℎ)𝑞𝑆𝑟𝑒𝑓𝑐̅ + 2𝑇m𝛿Tmz𝑙𝑇𝑚𝑥 (42)

𝛿𝑒 = −(1 − 𝜂)𝑅𝐻𝑀𝑆

𝐶𝑚𝛿𝑒𝑞𝑆𝑟𝑒𝑓𝑐̅

(43)

𝛿𝑇3 = 𝜂𝑅𝐻𝑀𝑆

𝑇a𝑙𝑇𝑎𝑧 (44)

𝜇𝑚 = 𝑡𝑎𝑛−1 𝛿Tmz

𝛿Tmx (45)

𝛿T1 = 𝛿T2 = 𝛿Tm = √𝛿Tmx2 + 𝛿Tmz

2 (46)

Figure (4) shows the main ducts’ tilt angle with respect to velocity in order to sustain the trim conditions. In Figure

(4), it can be seen that the aerodynamic forces become dominant after flight velocity 15m/s which is nearly stall speed.

Figure (5) shows the main and aft engine thrust ratios, and the elevator deflection with respect to flight velocity.

Table 8. Transition Flight Trim Values for Various Flight Velocities

𝑽∞ 𝝁𝒎 𝜹𝑻𝟏 𝜹𝑻𝟐 𝜹𝑻𝟑 𝜹𝒆

0.100 85.731 0.774 0.774 0.595 0.000

1.000 85.732 0.784 0.784 0.607 0.000

10.000 84.182 0.697 0.697 0.448 0.000

11.000 83.862 0.680 0.680 0.323 4.611

12.000 83.434 0.655 0.655 0.208 6.460

13.000 82.831 0.619 0.619 0.110 6.318

14.000 81.956 0.570 0.570 0.036 4.662

15.000 80.620 0.506 0.506 0.000 1.792

16.000 78.452 0.426 0.426 0.000 -1.706

19.500 31.933 0.114 0.114 0.000 -10.021

19.600 26.481 0.109 0.109 0.000 -10.195

19.700 20.454 0.104 0.104 0.000 -10.367

19.800 13.910 0.101 0.101 0.000 -10.536

19.900 6.971 0.099 0.099 0.000 -10.703


11

C. Longitudinal Stability Analysis For the longitudinal stability analysis, the linearized equations of motion are rewritten in the state space matrix

format which are shown in Eq. (47) - (49)38. The eigenvalues of the system matrix A have been determined for all

three flight mode around equilibrium point. According to the eigenvalues’ location, the stability investigation has been

performed.

The state space representation of linearized EoM;

[

∆�̇�∆�̇�∆�̇�

∆�̇�

] = [

𝑋𝑢

𝑍𝑢

𝑀𝑢 + 𝑀�̇�𝑍𝑢

0

𝑋𝑤

𝑍𝑤

𝑀𝑤 + 𝑀�̇�𝑍𝑤

0

0𝑢0

𝑀𝑞 + 𝑀�̇�𝑢0

1

−𝑔000

] [

∆u∆w∆𝑞∆𝜃

] + [

𝑋𝛿𝑒

𝑍𝛿𝑒

𝑀𝛿𝑒 + 𝑀�̇�𝑍𝛿𝑒

0

𝑋𝛿𝑡

𝑍𝛿𝑡

𝑀𝛿𝑒 + 𝑀�̇�𝑍𝛿𝑡

0

] [∆𝛿𝑒∆𝛿𝑡

] (47)

𝑨 = [

𝑋𝑢

𝑍𝑢

𝑀𝑢 + 𝑀�̇�𝑍𝑢

0

𝑋𝑤

𝑍𝑤

𝑀𝑤 + 𝑀�̇�𝑍𝑤

0

0𝑢0

𝑀𝑞 + 𝑀�̇�𝑢0

1

−𝑔000

] (48)

𝑩 = [

𝑋𝛿𝑒

𝑍𝛿𝑒

𝑀𝛿𝑒 + 𝑀�̇�𝑍𝛿𝑒

0

𝑋𝛿𝑡

𝑍𝛿𝑡

𝑀𝛿𝑒 + 𝑀�̇�𝑍𝛿𝑡

0

] (49)

Figure 3. Transition Flight Main Ducts’ Tilt Angle vs Flight Velocity

0 5 10 15 200

10

20

30

40

50

60

70

80

90

VFS

, m/s

, de

g

Main Ducts Tilting Angle wrt VFS

Figure 4. Transition Flight Ducts’ Tilt Angle vs Flight Velocity a) Main Ducts b) Aft Duct

0 5 10 15 200

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

Main Duct Thrust Ratio wrt VFS

VFS

, m/s

m

0 2 4 6 8 10 12 14 16 18 200

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

Aft Duct Thrust Ratio and Elevator Deflection wrt VFS

VFS

, m/s

3

0 2 4 6 8 10 12 14 16 18 20-20

0

20

e

3 : Aft Duct Thrust Ratio

e : Elevator Deflection


12

The eigenvalues of A matrixes for three flight phases are calculated and the following roots are obtained.

For hover flight condition;

𝑨𝒉𝒐𝒗𝒆𝒓 = [

0000

0000

0001

−𝑔000

] (50)

The eigenvalues of 𝑨𝒉𝒐𝒗𝒆𝒓 are;

𝜆1,2,3,4 = 0

Since all roots placed on the imaginary axis the system is unstable and stability/control augmentation system is

required for steady flight.

For cruise flight condition;

𝑨𝒄𝒓𝒖𝒊𝒔𝒆 = [

−0.07234−1,04960,011536

0

−0,16599−5,0637−5,2873

0

020

−2,01011

−9,81000

] (51)

The eigenvalues of 𝑨𝒄𝒓𝒖𝒊𝒔𝒆 for various flight conditions have been tabulated in Table 9.

Since all roots placed on the left hind side of the imaginary axis, the system is stable for cruise flight for the above

conditions.

For the transition flight mode, 20m/s, 15m/s, 10m/s, 5m/s, 0.1m/s flight velocities have been chosen to analysis

the stability behavior during transition. The roots of the system matrix 𝑨𝒕𝒓𝒂𝒏𝒔𝒊𝒕𝒊𝒐𝒏 have been plotted in Figure 5 and

tabulated in Table 10. It is determined from the calculations that the aircraft is for velocities greater that 13m/s. For

the velocities less than 13m/s the aircraft does not have inherent stability.


Flight Conditions Short Period Long (Phugoid) Period

𝑉∞ = 16 𝑚/𝑠; ℎ = 50 𝑚 −2,8483 ± 8,1056𝑖 −0,010212 ± 0,69056𝑖

𝑉∞ = 18 𝑚/𝑠; ℎ = 50 𝑚 −3,2001 ± 9,1243𝑖 −0,015711 ± 0,69018

𝑉∞ = 20 𝑚/𝑠; ℎ = 50 𝑚 −3,5523 ± 10,143𝑖 −0,020796 ± 0,68985𝑖

Figure 5. Pole Locations for Transition Flight a) All Poles b) Long Period Mode Poles

-4 -3 -2 -1 0 1

-10

-5

0

5

10

Pole Locations for Transition Flight

Real Axis

Ima

gin

ary

Axis

: VFS

= 20 m/s

: VFS

= 15m/s

: VFS

= 10m/s

: VFS

= 5m/s

: VFS

= 0.1 m/s

-0.1 -0.05 0 0.05 0.1 0.15 0.2 0.25-1.5

-1

-0.5

0

0.5

1

1.5Pole Locations for Transition Flight (Long Period Mode)

Real Axis

Ima

gin

ary

Axis

: VFS

= 20 m/s

: VFS

= 15m/s

: VFS

= 10m/s

: VFS

= 5m/s

: VFS

= 0.1 m/s


13

IV. Control

A. Control Strategy and Control Algorithm

The aircraft has three flight modes namely, vertical flight mode, transition flight mode and conventional flight

mode. Each flight mode requires individual control strategy. In the vertical flight mode, aircraft will be able to fly

toward all directions in X, Y, and Z axes. This will be achieved by small roll and pitch angle changings. By changing

the main ducts’ thrust individually, the roll attitude will be changed and the aircraft can fly in ±y direction. By

changing the aft duct thrust value, the pitch angle will be changed and aircraft can fly in ±x direction. The yaw angle

can be changed by increasing or decreasing the counter rotating propeller angular speeds. For the yaw angle the rudder

placed at the rear duct exit vane can be also used. The thrust differences will be obtained via RPM change of propeller.

In conventional flight mode, aileron, elevator and rudder will be used to control the roll, pitch and yaw attitude of the

aircraft, respectively. In transition flight mode, aerodynamic and thrust forces and moments will be dominant with

respect to velocity while the main ducts are tilting. The required rear propeller’s thrust change and the elevator

deflection have been summarized in Figure (4). For the transition phase, a control law, in order to get a smooth and

controlled flight behavior, has been developed and implemented in to a flight computer. In Table 11 the control

characteristics are summarized for three flight modes.

The PI/PID control method implemented to the displacement auto pilot system36 is chosen because of the ease of

modeling and coding. The displacement auto pilot included PID bloc diagram has been shown in Figure (6).


Flight Conditions Short Period Long (Phugoid) Period

𝑉∞ = 20 𝑚/𝑠; ℎ = 50 𝑚 −3,5523 ± 10,143𝑖 −0,020796 ± 0,68985𝑖

𝑉∞ = 15 𝑚/𝑠; ℎ = 50 𝑚 −2,6726 ± 7,596𝑖 −0,007252 ± 0,6907𝑖

𝑉∞ = 10 𝑚/𝑠; ℎ = 50 𝑚 −1,798 ± 5,043𝑖 0,0115 ± 0,6925𝑖

𝑉∞ = 5 𝑚/𝑠; ℎ = 50 𝑚 −0,9491 ± 2,467𝑖 0,0558 ± 0,699𝑖

𝑉∞ = 0,1 𝑚/𝑠; ℎ = 50 𝑚 0,1249 ± 0,142𝑖 −0,1428 ± 0,135𝑖

Table 11. Control Forces for Flight Modes

Flight Mode Pitch Control Roll Control Yaw Control

Vertical Flight Mode Aft duct thrust

adjustment

Main ducts thrust

adjustment, individually

Propeller angular speed

change,

Rudder in the downwash

of the aft duct

Transition Flight Mode Mixing Rudder in the downwash

of the aft duct

Conventional Flight Mode Elevator Aileron Rudder


14

In order to control the aircraft experimentally, a control code has been created separately. This code, basically,

consists of four subroutines which are initialization module, calculation module, sensor module and command receiver

module. The code is running in a micro-controller unit (MCU) onboard. In initialization module, the variables and the

port numbers of the MCU has been defined. In sensor module, the sensor data coming from gyro, accelerometer,

sonar, and compass are collected and filtered. The command receiver module receives the command signals coming

from transmitter and rearrange them as reference inputs. Finally, the calculation module gather the all sensor data and

control commands, then calculates the signal which will be sent to the motors and control surfaces’ actuators. The

main duty of the code is gathering the control inputs from radio transmitter, comparing them with the actual attitude,

calculating the command signals and sending them to the to the actuators and motors. The flowchart of the code is

given in Fig.7.

V. Flight Demonstration

A. Construction Studies

In order to test the control code, an indoor and outdoor

test frames have been constructed. The indoor test frame

has three motor arm, and at the end of each arm, there is a

fixed engine and a propeller. All three engines’ thrust

vectors point to the upward direction. A test bed, having a

rotational freedom around both x axis and y axis was also

constructed for the indoor tests. The indoor test frame and

test bed is shown in Fig. 8.

After experienced the indoor test frame and verified

the control code, an outdoor test frame has been

fabricated. The second model has two main propeller

engines without ducts at each tip of the wings and a third

engine on the tail boom. The front engines’ mounts

fabricated as they can tilt 0 to 95 degrees about y axis of

the aircraft. The third engine has also capability of rotating

30 degree in both direction perpendicular to tail boom in

Figure 6. Displacement autopilot include PID block diagram

Start InitializationAC ResponseBeginning of

the Loop

Receiving Pilot

Command

Gathering Sensor

Raw Data Calculating the Signals for

Motors and Control Surfaces

Sending the Control Signals

Arranging Reference

Inputs

Filtering the Raw

Data

Figure 7. Flowchart of control algorithm

Figure 8. Hover flight experiments with test frame

and rest bed.


15

order to control the yaw direction of the aircraft especially during hover flight. The fabricated outdoor test frame is

shown in Fig. 9. With the outdoor test frame, the vertical take-off, hover flight and vertical landing have been achieved

successfully. Due to the lack of ducts and inappropriate aerodynamic design the transition flight was not achieved.

Currently, construction of the prototype aircraft has been started and some parts have been manufactured which

are shown in Fig. 10. It is accepted that the final assembly of the prototype aircraft will be completed in December

and immediately after that the flight test will start.

VI. Conclusion

In this study, a tilt duct VTOL UAV was designed. The stability and control characteristics have been investigated

theoretically. The main feature of the design is the tiltable, ducted, counter rotating propeller placed at the wing tips

and between tail booms. The control calculations showed that the aircraft is stable for cruise flight for velocities greater

than 13m/s. On the other hand, the aircraft does not have inherent stability for hover and flight velocities less than

13m/s. A control strategy approach and a control algorithm with respect to this approach have been developed. The

control algorithm has been tested successfully with low cost indoor and outdoor test frames which are designed and

fabricated for this purpose. After initial indoor and outdoor flight tests, the construction of the first prototype of the

final aircraft has been started. The construction and the flight test are planning to be finished in Dec 2015.

Acknowledgments

The authors acknowledge the financial support provided by The Scientific and Technological Research Council of

Turkey (TUBITAK) under grant 213M344, BOEING Executive Focal, and Anadolu University Scientific Research

Projects Department under grant 1308F310.

Figure 9. Hover flight experiments with outdoor test frame.

Figure 10. The prototype aircraft construction a) Fuselage b) Wing assembly c) Main duct quarter


16

References 1Franklin, J. A., Dynamics, Control, and Flying Qualities of V/STOL Aircraft, 1st ed., AIAA Education Series, Virginia, 2002,

pp. xi, 9-11. 2Maisel, M. D., Giulianetti, D. J., Dugan, D. C., The History of the XV-15 Tilt Rotor Research Aircraft: From Concept to Flight,

Monographs in Aerospace History #17, the NASA History Series, Washington, D.C., 2000. 3Tesla, N., New York, U.S. Patent Application for a “Method of Aerial Transportation,” Serial No: 499,519, filed 09 Sep. 1921. 4Piccirillo, A. C., German V/STOL Fighter Program: A Quest for Survivability in a Theater Nuclear Environment, AIAA,

Virginia, 1997, pp. xi-xiii 5Austin, R., Unmanned Aircraft Systems: UAVS Design Development and Deployment, 1st ed., Wiley & Sons, UK, 2010, pp.7. 6 Jeong, J., Yoon, S., Kim, S., Suk, J., “Dynamic Modeling and Analysis of a Single Tilt-Wing Unmanned Aerial Vehicle”,

53rd AIAA Aerospace Sciences Meeting, AIAA 2015-1804, Kissimmee, Florida, 2015. 7Bell Helicopter, Bell Eagle Eye Pocket Guide. URL:http://epic.org/privacy/surveillance/spotlight/0805/eagle.pdf [cited

10 May 2013]. 8Choi, S., Kang, Y., Chang, S., Koo, S., Kim, J.M., “Development and Conversion Flight Test of a Small Tiltrotor Unmanned

Aerial Vehicle,” Journal of Aircraft, Vol. 47, No. 2, March–April 2010, pp. 730-732. 9Song, Y., and Wang, H., “Design of Flight Control System for a Small Unmanned Tilt Rotor Aircraft Longitudinal Flight

Dynamic Analysis of an Agile UAV”, Chinese Journal of Aeronautics, Vol. 22, 2009, pp. 250-256. 10Israel Aerospace Industry, Tactical VTOL UAVs. URL:http://www.iai.co.il/2013/36719-

en/BusinessAreas_UnmannedAirSystems.aspx [cited 1 Nov 2015]. 11 Özdemir, U., Aktaş, Y. O., Vuruşkan, A., Dereli, Y., Tarhan, A. F., Demirbağ, K., Erdem,A., Kalaycioğlu, G. D., Özkol, İ.,

İnalhan, G., “Design of a Commercial Hybrid VTOL UAV System”, Journal of Intelligent & Robotic Systems, Vol. 74, No. 1,

2014, pp. 371-393 12RHoneywell Aerospace, Military Avionics. URL: https://aerospace.honeywell.com/products/military-avionics/t-

hawk-mav [cited 01 Nov 2015] 13Wikipedia The Free Encyclopedia, Honeywell RQ-16 T-Hawk. URL: https://en.wikipedia.org/wiki/Honeywell_RQ-

16_T-Hawk [cited 01 Nov 2015] 14Lipera, L., Colbourne, J. D., Tischler, M. B., Mansur, M. H., Rotkowitz, M. C., Patangui, P., “The Micro Craft iSTAR Micro

Air Vehicle: Control System Design and Testing,” American Helicopter Society 57th Annual Forum, Washington, DC, 2001. 15Johnson, E. N., Turbe, M. A., “Modeling, Control, and Flight Testing of a Small Ducted-Fan Aircraft” Journal of Guidance,

Control, and Dynamics, Vol. 29, No. 4, 2006, pp.769-780. 16Ohanian, O. J., Gelhausen, P. A. and Inman, D. J., “A Compact Method for Modeling the Aerodynamics of Ducted Fan

Vehicles,” 48th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition, 4 - 7 January

2010, Orlando, Florida. 17Çetinsoy. E., Sırımoğlu., E., Öner, T. K., Hançer, C., Ünel, M., Akşit, M. F., Kandemir, İlyas, Gülez, K., “Design and

development of a tilt-wing UAV”, Turkish Journal of Electrical Engineering and Computer Sciences, Vol.19, No.5, 2011, pp. 733-

741. 18 Öner, K.T., Çetinsoy, E., Sırımoğlu, E., Hançer, C., Ünel, M., Akşit, M. F., Gülez, K., Kandemir, İ., “Mathematical modeling

and vertical flight control of a tilt-wing UAV”, Turkish Journal of Electrical Engineering and Computer Sciences, Vol.20, No.1,

2012, pp. 149-157. 19Sato, M., Muraoka, K., “Flight Controller Design and Demonstration of Quad-Tilt-Wing Unmanned Aerial Vehicle” Journal

of Guidance, Control, And Dynamics, Vol. 38, No. 6, pp. 1071-1082. 20Suziki, S., Zhijia, R., Hor,ta, Y., Nonami, K., K,mura, G., Bando, T., Hirabayashi, D., Furuya, M., Yasuda, K., “Attitude

Control of Quad Rotors QTW-UAV with Tilt Wing Mechanism”, Journal of System Design and Dynamics, Vol.4, No.3, 2010, pp.

416-428. 21Totoki, H., Ochi, Y., Sato, M., Muraoka, K., “Design of Gain Scheduled Stability and Control Augmentation System for

Quad-Tilt-Wing UAV”, 53rd AIAA Aerospace Sciences Meeting, AIAA 2015-1804, Kissimmee, Florida, 2015. 22McCormick, B. W., Aerodynamics of V/STOL Flight, 1st ed., Academic Press, Florida, 1967, pp. 252. 23Abrego, A.I., Bulaga, R. W., “Performance Study of a Ducted Fan System”, American Helicopter Society Aerodynamics,

Acoustics, and Test and Evaluation Technical Specialists Meeting, San Francisco, CA, 2002. 24Koç, S. T., Yılmaz S., Erdem D., Kavsaoğlu M. Ş., “Experimental Investigation of a Ducted Propeller”, 4th European

Conference for Aerospace Sciences (EUCASS 2011), Saint Petersburg, 2011. 25Bo, W., Zheng, G., Peng, W., Shangqiu, S., Zhongxi, H., “Investigation of the Aerodynamic Characters of Ducted Fan

System”, World Academy of Science, Engineering and Technology, Vol. 6, No. 9, 2012, pp. 1178-1181 26Colin, P.E., Williams, J., “The Aerodynamics of V/STOL Aircraft,” Lecture Notes, AGARD AD688921, 1968. 27McCormick, B. W., Aerodynamics of V/STOL Flight,1st ed., Academic Press, Florida, 1967, pp. 252. 28Armutcuoğlu, Ö., Kavsaoğlu, M. Ş., and Tekinalp, O., “Tilt Duct Vertical Takeoff and Landing Uninhabited Aerial Vehicle

Concept Design Study,” Journal of Aircraft, Vol. 41, No. 2, March–April 2004, pp. 215-223. 29Kavsaoğlu, M. Ş., Öznalbant, Z., “Sabit Kanatlı DİK İHA Benzetim ve Kontrolü,” 7nci Havacılık Sempozyumu, Kayseri,

2008, pp. 99-100 30Öznalbant, Z., Kavsaoğlu, M. Ş., “Design and Flight Test Study of a VTOL UAV”, 53rd AIAA Aerospace Sciences Meeting,

AIAA 2015-1903, Kissimmee, Florida, 2015.

http://epic.org/privacy/surveillance/spotlight/0805/eagle.pdf

http://www.iai.co.il/2013/36719-en/BusinessAreas_UnmannedAirSystems.aspx

http://www.iai.co.il/2013/36719-en/BusinessAreas_UnmannedAirSystems.aspx

https://aerospace.honeywell.com/products/military-avionics/t-hawk-mav

https://aerospace.honeywell.com/products/military-avionics/t-hawk-mav

https://en.wikipedia.org/wiki/Honeywell_RQ-16_T-Hawk

https://en.wikipedia.org/wiki/Honeywell_RQ-16_T-Hawk


17

31Raymer, D. P., Aircraft Design: A Conceptual Approach, AIAA Education Series, Virginia, 2006, Chap. 20. 32Etkin, B., and Reid, L.D., 1996. Dynamics of Flight Stability and Control, 3rd ed., John Wiley and Sons Press, Toronto, 1994,

Chap. 4. 33Yechout, T. R., Morris, S. L., Bossert, D. E., and Hallgre W. F., Introduction to Aircraft Mechanics Perfomance Static

Stability Dynamic Stability and Classical Feedback Control, AIAA Education Series, Virginia, 2003, Chap. 4. 34Roskam, J., Airplane Flight Dynamics and Automatic Flight Controls, Roskam Aviation and Engineering Corporation,

Kansas, 1979, Chap. 4. 35Kavsaoğlu, M.Ş., “Advanced Flight Dynamics Lecture Notes” Istanbul Technical University, Lecture Notes, 2005. 36Stevens, B.L., Lewis, F.L., “Aircraft Control and Simulation”, 2nd ed., John Wiley and Sons Press, New Jersey, 2003, pp.117-

119. 37Chapra, S.C., Canale R. P., (Trans.: Heperkan, H., Kesgin, U.), Numerical Methods for Engineers (Trans.: Mühendisler için

Sayısal Yöntemler), Literatür Yay., Vol. 4, 2003, İstanbul. 38Nelson, R.C., Flight Stability and Automatic Control, 2nd ed. McGraw Hill Education (Special Indian Edition), New Delphi,

2007, pp. 149-150, 292-302

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