FlamingoDetailed Design of an Emergency Relief UAV

Principal TutorP. Roling

CoachesM. Alharbi

H. Chen

Design Synthesis Exercise Group 6

Rene van den Berg 1515403 Jonas Laeret 1396374Rinze Bruining 1506773 Pepijn Meeuwissen 1534505Joep van Genuchten 1369148 Olivier Mulder 1236393Darsini Kathirgamanathan 1509497 M.A.C. Perera 1551019Ayesha Khan 1543970 Robin Vermeij 1357263

June 28, 2011

ii Nomenclature

Preface

This report is the fourth and final progress report of Design Synthesis Exercise Group 6. It describesthe results of the design process that led to the design of an Unmanned Aerial Vehicle (UAV) foremergency relief purposes. The emphasis of the report lies on the design philosophy, the used method-ologies and the results obtained during the detailed design phase. The main fields of investigation arethe aerodynamics, the structures and the simulation of the UAV.

Readers who are particularly interested in aerodynamic, stability and performance aspects of thedesigned UAV should direct their attention to chapter 3. Chapter 4 focuses on an in depth analysisof various structures of the UAV, whereas chapter 5 deals with the numerical simulation modelingthe behavior of the aircraft while it is performing the vertical attitude take off and landing (VATOL)maneuver. Chapters 6 and 7 illustrate the subsystems of the UAV and the operations and logisticsrespectively.

We would like to express our gratitude to our tutor P. Roling for his continuous guidance and inputduring the Design Synthesis Exercise. The same holds for our coaches M. Alharbi and H. Chen. Wealso would like to thank M. Abdalla, C. Kassapoglou and J. Hol for their help with the structural de-sign and analysis. Furthermore, we want to thank M.D. Pavel for her help on the MATLAB programused in the design of the propellers. Additionally, we would like to thank M. Voskuijl for his inputon the tilt rotor concept design. We are also grateful for the advice given by our oral presentationlecturer B. van der Laaken. Finally, special thanks are given to E. Mooij. We greatly appreciate hiscontributions to the numerical simulation of the VATOL maneuver. Without his help, the simulationwould not have reached the level that it has now.

iii

iv 0 Preface

Summary

The efficient delivery of emergency humanitarian aid can drastically minimize the aftermath of anatural disaster. This calls upon a system of great flexibility that can easily adapt to variableoperating environments. The unpredictability associated with natural disasters often dismisses thepossibility of using conventional means of transport to achieve this objective. Hence, delivery by airbecomes one of the more feasible options creating a demand that is currently met by helicopters.

Helicopter’s enhanced control and its ability to hover, drastically simplifies the delivery process.However, due to inefficiencies in their use with respect to sustainability and fuel consumption, Un-manned Aerial Vehicles (UAV) can be seen as a viable substitute in the near future.

This particular Design Synthesis Exercise (DSE) project deals with the design of a UAV to pro-vide humanitarian aid to the residents of areas struck by natural disasters. The UAV should be ableto perform the mission in an efficient, cheap and environmentally friendly manner with respect to ahelicopter.

The analysis of customer and designer based requirements and the exploration of available designoptions led to the preliminary design of three different concepts. These concepts, a tilt rotor concept,helicopter concept and a tail sitter fixed wing concept, were analyzed in detail. The trade off per-formed yielded the tail sitter concept as the winning concept that has been designed in detail. Thetail sitter from then on was designated the ’Flamingo’.

The Flamingo UAV is a unique, novel design, which incorporates the attributes of a blendedwing body with a canard type configuration. It includes two wing mounted propellers driven by asingle Wankel engine, Mistral G-200. Flamingo accomplishes the requirement of landing and takingoff vertically by performing a pitch up maneuver. In other words, the UAV pitches up steeply,transferring the lifting load from its wings to its propellers. It will then slowly descent, where it willland on its payload. After the releasing of the payload, the UAV will then take off vertically. Thisinherently implies that its thrust to weight ratio is larger than one, after the release of the payload.

The UAV also meets the 75% carbondioxide emission reduction requirement with respect to aconventional helicopter. This illustrates the sustainable nature of the UAV. Furthermore, dependingon the payload weight of the aircraft, the design of the Flamingo allows for total world coverage.

The tail sitter concept can be modeled as a conventional fixed wing aircraft during its horizontalflight regime and as a rotor craft while it is performing the latter part of the pitch up maneuver. Thishas a direct consequence for the design of the propellers, in the sense that it is a compromise betweena helicopter rotor and a conventional aircraft type propeller. Blade element momentum theory is usedto compute the thrust and power delivered by the propeller. Furthermore, the induced air velocity bythe propeller is determined using a numerical simulation. The determination of this induced velocityis vital for modeling the behavior of the aircraft while it is performing the VATOL maneuver, as thisbecomes the only airflow over the control surfaces allowing for pitch, yaw and roll control in no windconditions.

v

vi 0 Summary

A detailed aerodynamic analysis of this novel concept is performed using vortex lattice method (VLM)using the MATLAB based program TORNADO. TORNADO models the main wing, canard and thevertical fins of the UAV. The wing blended fuselage is modeled using the NACA 0025 airfoil. Thisprogram outputs the lift distribution of both the wing and the canard. The drag estimation wascomputed using both TORNADO and the XFOIL based program XFLR5. The drag coefficients ofthe cargo pod is computed using XFLR5 and the model in TORNADO gives the drag coefficient ofthe blended wing body. This output has been corrected using general aerodynamic relation to depicta more realistic picture.

The main wing is placed as far back as possible with respect to the aircraft, in order to shiftthe aerodynamic center back and increasing the stability margin. The stability of the UAV duringconventional and hovering flight with and without payload, is improved by iterating between variouscanard sizes and locations. The result is a neutrally statically stable aircraft with payload. Withoutpayload, the aircraft has a positive stability margin and is therefore statically stable. But moreimportantly, the UAV is proved to be controllable during its most critical and challenging flight phaseof the mission, the hovering phase.

All flight derivatives and control deflection derivatives that form the input to the simulation ofthe VATOL maneuver are computed using TORNADO.

An in depth structural analysis has also been performed on the most critical parts of the UAV.This analysis was also coupled with finite element modeling using the Generative Structural Analysisworkbench in CATIA.

The internal structure of the wing box is designed in detail to account for numerous possiblefailure modes and load types. A numerical simulation is utilized to compute the optimum stringerand rib pitch. The internal structure of the canard is also sized in a similar manner. The canardbooms are modeled to be clamped at their connection to the engine mount and have been sized forminimum deflections to ensure that the canard position is not hindered due to elastic deformation.Furthermore, the main landing gear is sized for hard landing conditions, where the UAV lands on asingle wheel with a landing gear load factor of 2. The tail landing gear is sized for the maximumstatic load acting on the tail wheel and steers the aircraft while grounded. The main landing gearconsists of a main strut and two side struts. The structure is modeled by one roller and two hingesupports. The landing gear retracts parallel to the two booms connecting the canard to the aircraft.The cargo pod is designed to have a payload volume of 1m3. It is made of white spruce strengthenedby stiffeners to withstand the loads during the landing. The cargo pod is made from a biodegradablematerial, reinforcing the sustainable benefits of this design. The design of the shafts and gearboxestransferring the power from the engine to the propellers are also analyzed. The gearboxes use a fixedgear ratio, worm gears and bevel gears are used to connect one shaft to another.

The VATOL maneuver performed by the UAV is a critical aspect of the design. This procedureis modeled using two simulations in MATLAB and Simulink environments. Two simulations weremade such that one can be verified with the other. However, the MATLAB simulation failed to achieveaccurate results within the available time period. The euler angles suffer from Gimbal lock while theUAV is in vertical attitude. Hence, the equations of motion have been determined in quaternion form.These are used to create a state space system.

Both the simulation utilizes the same Linear Quadratic Regulator (LQR) controller, which gener-ates the gain needed to control the aircraft during the VATOL maneuver. The results of the Simulinksimulation illustrates that the designed UAV can indeed perform the VATOL maneuver, while lim-iting the descent acceleration to below 2g. Furthermore, it confirms that no stability problems arefaced during the vertical take off, after the cargo is released and thus the center of gravity shifted.The control surfaces were confirmed to be sufficiently effective to maintain the stability of the aircraftduring the VATOL phase under the propeller induced velocity conditions.

0 Summary vii

The Flamingo contains a triple redundant autopilot system specifically designed for UAVs and in-corporates a long distance landing sensor for landing site scanning purposes and a short distancelanding sensor. These sensors function during both day and night increasing the flexibility of theUAV. Flamingo’s communication system bases itself on the use of broadband satellite Internet. Re-dundancies were considered when designing the fuel system of the Flamingo. Its fuel tank is especiallydesigned to minimize sloshing, as this can lead to stability issues while the VATOL maneuver is per-formed.

The maximum take off weight of the flamingo is calculated to be 7140N . It has a wing span of7.45m. The Flamingo’s stall speed and maximum speed are 28m/s and 77m/s respectively. With acruise speed of 55m/s, it is able to perform six missions per day. When the UAV is fully loaded therange is 590km which is in general large enough for emergency relief operations. During a one week-operation, the UAS consisting of ten UAVs can provide 20, 000 beneficiaries with 5kg of humanitarianaid.

Flamingo was developed to very high level of detail. Mainly constrained by time, further analy-sis of the structural parts using finite element models were not possible. Numerous failure modes ofstructural parts still need to be analyzed, namely for buckling, fatigue and vibrational loads.

The cargo pod should be designed in much greater detail with respect to aerodynamics. Currently,only the structural attributes of this is considered. The simulation confirms that the VATOL analysiscan be performed by the Flamingo. The next step is to design an auto pilot for this maneuver basedon the simulation.

viii 0 Summary

Table of Contents

Preface iii

Summary v

List of Symbols xiii

1 Introduction 1

2 Design Aspects 3

2.1 Project objective . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2.2 Emergency relief UAV requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.2.1 Customer requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.2.2 Designer imposed requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.3 Considered concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.3.1 The helicopter concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.3.2 The tiltrotor concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

2.3.3 The tailsitter concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

2.3.4 Trade-off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

2.4 Features of the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.4.1 Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.4.2 VATOL capability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.4.3 Cargo pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.4.4 Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.4.5 Flexibility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.5 Mission profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.5.1 Flight mission . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.6 Sustainable development . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.6.1 Manufacturing the UAV . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.6.2 Operating the UAV . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.7 Market analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.7.1 Disasters and customer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.7.2 Humanitarian relief . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

2.7.3 Positioning the product . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

ix

x Contents

3 Aerodynamics, Stability and Performance 15

3.1 Aerodynamic analysis methodology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.1.1 The vortex lattice method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3.1.2 Implementation in the simulation . . . . . . . . . . . . . . . . . . . . . . . . . . 16

3.1.3 The aerodynamic model of the Flamingo . . . . . . . . . . . . . . . . . . . . . . 16

3.2 Analysis of the aerodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

3.2.1 CL − α and CL − CD curves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

3.2.2 Main wing lift distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

3.3 Static stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

3.3.1 The stability margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

3.3.2 Static stability during horizontal flight . . . . . . . . . . . . . . . . . . . . . . . 19

3.3.3 Static stability during hover flight . . . . . . . . . . . . . . . . . . . . . . . . . 19

3.4 Dynamic stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

3.5 Propeller design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

3.6 Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

3.6.1 Performance diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

3.6.2 Thrust to weight ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

3.6.3 Range and endurance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

3.6.4 Rate of climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

3.6.5 Stall speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

3.6.6 Gliding rate of descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

3.6.7 Take off field length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

3.7 Mission performance analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

3.8 Overview of aerodynamic and performance parameters . . . . . . . . . . . . . . . . . . 27

4 Structural Design & Analysis 29

4.1 Wing box . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

4.1.1 Method used to determine normal stresses and shear flows . . . . . . . . . . . . 29

4.1.2 Loads and stresses acting on each cross section . . . . . . . . . . . . . . . . . . 31

4.1.3 Rib and stringer spacing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

4.1.4 Finite element analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.1.5 Material and weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38

4.2 Canard design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38

4.2.1 Shear stress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

4.2.2 Stiffeners of the canard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.2.3 Spar flanges of the canard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

4.2.4 Material and weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

4.3 Cargo pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42

4.3.1 Centre of gravity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.3.2 Stress analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.3.3 Set up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4.3.4 Material selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

4.4 Cargo release mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

4.5 Landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

4.5.1 Shock absorber . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.5.2 Main landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.5.3 Tail landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49

4.6 Booms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

4.6.1 Stress analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

4.6.2 Material selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

4.6.3 Boom connection to engine mount . . . . . . . . . . . . . . . . . . . . . . . . . 53

Contents xi

4.6.4 Finite element analysis on booms . . . . . . . . . . . . . . . . . . . . . . . . . . 53

4.7 Control surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54

4.7.1 Attaching the flaperons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

4.7.2 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

4.8 Shafts and gear boxes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

5 VATOL Simulation 57

5.1 Flight dynamics and quaternions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57

5.2 Derivation of the linear system of equations . . . . . . . . . . . . . . . . . . . . . . . . 57

5.3 Aircraft controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

5.4 Aircraft guidance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

5.5 Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61

5.5.1 Assumptions and aerodynamic data . . . . . . . . . . . . . . . . . . . . . . . . 62

5.5.2 Simulink implementation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62

5.6 Results of the simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

5.6.1 Pitch-up maneuver . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

5.6.2 Vertical take off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

6 Subsystem Design 69

6.1 Power system design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 69

6.2 Electrical block diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

6.3 Control & navigation system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.3.1 Autopilot chip set . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.3.2 Short distance landing sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.3.3 Long distance landing sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.4 Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

6.5 Data handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

6.6 Support system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

7 Operations, Logistics & System Engineering 75

7.1 Operations & logistics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

7.1.1 Transportation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

7.1.2 Personnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

7.1.3 Material handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

7.2 Production plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

7.2.1 Manufacturing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

7.2.2 On site assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

7.3 Cost break-down structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

7.3.1 Unit production cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

7.3.2 Direct operating cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

7.4 Operational benefit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

7.5 Risk map . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

7.6 RAMS analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80

7.6.1 Reliability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80

7.6.2 Availability and maintainability . . . . . . . . . . . . . . . . . . . . . . . . . . . 81

7.6.3 Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81

7.7 Functional flow diagram & functional breakdown structure . . . . . . . . . . . . . . . . 81

7.8 Future project design and logic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82

8 Conclusions 87

xii Contents

9 Recommendations 899.1 Recommendations for the project in general . . . . . . . . . . . . . . . . . . . . . . . . 899.2 Recommendations for aerodynamics, stability and performance . . . . . . . . . . . . . 899.3 Recommendations for structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 909.4 Recommendations for the simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 909.5 Recommendations for the subsystems . . . . . . . . . . . . . . . . . . . . . . . . . . . 919.6 Recommendations for operations and logistics . . . . . . . . . . . . . . . . . . . . . . . 91

Appendices

A Technical drawings 97

B State Space System Matrix Inputs 103

C Compliance Matrix 105

List of Figures

2.1 Artist impression of the Helicopter concept . . . . . . . . . . . . . . . . . . . . . . . . 62.2 Artist impression of the Tilt Rotor concept . . . . . . . . . . . . . . . . . . . . . . . . 72.3 Artist impression of the Tailsitter concept . . . . . . . . . . . . . . . . . . . . . . . . . 72.4 Artist impression of the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

3.1 Aerodynamic model of the tailsitter using VLM in TORNADO . . . . . . . . . . . . . 163.2 Aerodynamic model of the cargo pod using XFLR5 . . . . . . . . . . . . . . . . . . . . 173.3 The CL − α curve of the tailsitter UAV . . . . . . . . . . . . . . . . . . . . . . . . . . 183.4 The CL − CD curve of the tailsitter UAV, with and without payload . . . . . . . . . . 183.5 The lift distribution of the main wing for α = 3◦ deg and V = 55m/s . . . . . . . . . . 193.6 The Cm − α curve of the tailsitter UAV with and without payload . . . . . . . . . . . 203.7 Free body diagram of the UAV in hovering phase . . . . . . . . . . . . . . . . . . . . . 203.8 The Cmδ,elevon curve of the tailsitter UAV with and without payload during hover . . . 213.9 Hovering pitch angle versus resulting force in Xearth-direction with and without payload 213.10 Angle of attack of the Flamingo during the phugoid motion . . . . . . . . . . . . . . . 213.11 Velocity of the Flamingo during the phugoid motion . . . . . . . . . . . . . . . . . . . 213.12 Delivered thrust, required power and induced velocity of a single proprotor for varying

pitch angles and velocity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233.13 Performance diagram for the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . 243.14 The thrust to weight ratio versus altitude with payload . . . . . . . . . . . . . . . . . 243.15 The maximum hovering altitude versus payload weight . . . . . . . . . . . . . . . . . . 243.16 The rate of climb versus altitude with and without payload . . . . . . . . . . . . . . . 253.17 The stall speed versus altitude with and without payload . . . . . . . . . . . . . . . . 253.18 V/a diagram during the take off run . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

4.1 Forces acting on the wing box and their locations . . . . . . . . . . . . . . . . . . . . . 304.2 Overview of the variables used for the rib and stringer sizing . . . . . . . . . . . . . . 324.3 Top view of the rib and stringer distribution on the bottom plate . . . . . . . . . . . . 334.4 Bottom view of the rib and stringer distribution on the top plate . . . . . . . . . . . . 334.5 Overview of the outer wing box model . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.6 Cross sections of the wing box . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.7 Stress distribution in the outer wing box using FEM analysis . . . . . . . . . . . . . . 354.8 Displacement of the outer wing box using FEM analysis . . . . . . . . . . . . . . . . . 364.9 General overview of the internal fuselage structure . . . . . . . . . . . . . . . . . . . . 374.10 Overview of the displacements of the inner wing box at the extreme flight conditions

case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

xiii

xiv List of Figures

4.11 The wing box cross-section of the canard . . . . . . . . . . . . . . . . . . . . . . . . . . 394.12 Moment diagram due to lift (left) and moment diagram due to drag (right) . . . . . . 394.13 The cross-section of the stiffener and the spar flange . . . . . . . . . . . . . . . . . . . 414.14 Cargo pod configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 424.15 Cargo pod fold lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 454.16 Release mechanism of the cargo pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . 464.17 Sketch of the main landing gear and cross-section of the beam . . . . . . . . . . . . . . 494.18 The tail gear location (left) and the cross-section of the beam (right) . . . . . . . . . . 504.19 Stress Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524.20 Boom stress distribution due to loads encountered during the cruise flight regime . . . 534.21 Translational displacements of the boom due to loads faced during cruise . . . . . . . 544.22 Boom stress distribution due to worst case scenario loads . . . . . . . . . . . . . . . . 544.23 Control surface design. a) Main design; b) Open section shear flows; c) Closed section with

additional circular shear flows; d) Flaperons support. . . . . . . . . . . . . . . . . . . . . . . . 554.24 Shaft and gearbox design. a) Rotations and RPMs; b) Gears and bearings. . . . . . . . . . . 56

5.1 The gains are computed at several instances when using gain scheduling . . . . . . . . 605.2 Flight Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 635.3 Simulink controller block . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 635.4 Flight Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 645.5 simulink guidance block . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 645.6 Altitude versus distance to target during vertical landing . . . . . . . . . . . . . . . . 655.7 Pitch angle versus time during the vertical landing . . . . . . . . . . . . . . . . . . . . 655.8 Velocity versus time during the vertical landing . . . . . . . . . . . . . . . . . . . . . . 665.9 Altitude versus horizontal distance from target during the vertical take off . . . . . . . 665.10 Pitch angle versus time during the the vertical take off . . . . . . . . . . . . . . . . . . 665.11 Velocity versus time during the vertical take off . . . . . . . . . . . . . . . . . . . . . . 67

6.1 Electrical Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 706.2 Satellite Internet coverage of the BGAN network . . . . . . . . . . . . . . . . . . . . . 726.3 Satellite Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 726.4 Data Handling Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73

7.1 A400M layout where ten UAVs are loaded with support staff and support equipment . 767.2 Functional Breakdown Structure of a typical emergency relief mission . . . . . . . . . 837.3 Functional Flow Diagram of a typical emergency relief mission . . . . . . . . . . . . . 847.4 Gantt chart for further development . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85

A.1 Isometric view of the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98A.2 Top view of the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99A.3 Front view of the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100A.4 Side view of the Flamingo without cargo pod . . . . . . . . . . . . . . . . . . . . . . . 101A.5 Side view of the Flamingo with cargo pod . . . . . . . . . . . . . . . . . . . . . . . . . 102

List of Symbols

Symbol Unit Meaning Symbol Unit Meaning

A m2 Area ~q vector Quaternion vector

Br m2 Boom area r rad/s Yaw rate

Ce − Dimensionless change ofquaternion input 1 Rc −

Ratio of actual and crit-ical stress: compressionbuckling

Cd m/s2 2D Drag coefficient Rs − Ratio of actual and criti-cal stress: shear buckling

CD m/s2 3D Drag coefficient St m Tire deflection

Cl m/s2 2D Lift coefficient Sx m3 Statical moment of areaaround x-axis

CL m/s2 3D Lift coefficient Sy m3 Statical moment of areaaround y-axis

Cp − Dimensionless ∆ p1 Sz m3 Statical moment of areaaround z-axis

Cq − Dimensionless ∆ q1 t m Thickness

Cr − Dimensionless ∆ r1 T N Thrust

d m Stiffener pitch td m Web thickness

E Pa Young’s modulus ~u vector Control input vector

e − Quaternion input V m/s Speed

e0 − Scalar quaternion input V N Shear force

Fnormal N Normal force Vs m/s Sink speed

xv

xvi 0 List of Symbols

Symbol Unit Meaning Symbol Unit Meaning

g m/s2 Gravitational accelera-tion parameter

W N Weight

H m Height x m Roll-axis

I m4 Moment of inertia ~x vector State vector

K − Column effective factor y m Pitch-axis

Kc − Buckling constant z m Yaw-axis

L N Lift α rad Angle of attack

L or l m Length γ rad Flight path angle

lav m Available length δ∆T− Change in ∆T

1

M Nm Moment δa rad Aileron deflection

N N Normal force δc rad Canard deflection

n − Number of stiffeners δe rad Elevator deflection

ns − Shock absorber efficiency δr rad Rudder deflection

nt − Efficiency of the tire δT − Thrust setting

P N Normal force θ rad Pitch angle

p rad/s Roll rate ν − Poisson’s ratio

Pcr N Critical buckling force κ m−1 Curvature

Q N Statical moment of area σ Pa Normal stress

q rad/s Pitch rate τ Pa Shear stress

qs N/m Shear flow τcr Pa Critical shear stress

Table 1: List of symbols

1An additional subscript indicates that the derivate of the parent is taken with respect to the child.

List of Tables

1 List of symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiv

2.1 Mission performance of the Helicopter configuration . . . . . . . . . . . . . . . . . . . 6

2.2 Mission performance of the Tilt Rotor configuration . . . . . . . . . . . . . . . . . . . 6

2.3 Mission performance of the TailSitter Canard configuration . . . . . . . . . . . . . . . 7

2.4 Trade off criteria and weights and scores for each concept . . . . . . . . . . . . . . . . 8

2.5 R-44 Mission profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2.6 Number of beneficiaries for a two day operation . . . . . . . . . . . . . . . . . . . . . . 13

2.7 Number of beneficiaries for a seven day operation . . . . . . . . . . . . . . . . . . . . . 13

2.8 The SWOT analysis of the Flamingo . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.1 Properties of the CL − α curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

3.2 Basic proprotor configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

3.3 Typical (single) proprotor settings and quantities . . . . . . . . . . . . . . . . . . . . . 22

3.4 The range and endurance of the UAV . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

3.5 The total time and fuel consumed for a 200km range, 3000m altitude mission . . . . . 27

3.6 Costs per component of a 200km range, 3000m altitude mission . . . . . . . . . . . . . 27

3.7 Overview of the aerodynamic parameters of the UAV . . . . . . . . . . . . . . . . . . . 28

3.8 Overview of the performance parameters of the Flamingo . . . . . . . . . . . . . . . . 28

4.1 Loads, moments, stresses and shear flows acting on the wing box . . . . . . . . . . . . 31

4.2 Rib Spacing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.3 Loads applied to the FEM of the outboard wing box . . . . . . . . . . . . . . . . . . . 35

4.4 Material properties of aluminum 7075 T76511 . . . . . . . . . . . . . . . . . . . . . . . 38

4.5 Bending moments, normal stress and normal force on booms at different section in thewing box of the canard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.6 Shear flow and the shear stresses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

4.7 Number of stiffeners needed on the top plate and the rear spar . . . . . . . . . . . . . 41

4.8 Dimensions of the legs of the pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.9 Stresses on the leg of the pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

4.10 Stresses on the pivot of the pod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4.11 Stiffeners needed for the plates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

4.12 White Spruce Picea Glauca longitudional properties . . . . . . . . . . . . . . . . . . . 45

4.13 The normal stress and the shear stress at different location in the landing gear . . . . 48

4.14 The stresses in the tail landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

4.15 Boom diameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51

xvii

xviii List of Tables

4.16 Stresses on the booms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 514.17 Stresses on the booms worst case . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524.18 Epoxy Carbon fibre (SMC) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524.19 Comparison of FEM with analytical solution . . . . . . . . . . . . . . . . . . . . . . . 534.20 Main sizing parameters, determined analytically . . . . . . . . . . . . . . . . . . . . . 554.21 Results of shaft and gear box structural analysis. . . . . . . . . . . . . . . . . . . . . 56

6.1 Different components and their masses . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

7.1 Flamingo unit cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 787.2 Direct operating cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 797.3 Risk map . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81

8.1 Flamingo characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88

C.1 Compliance Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 107

Chapter 1

Introduction

Natural disasters are known to leave devastation in their wake. The unpredictable nature of thesedisasters requires efficient ways to distribute humanitarian aid quickly to the sites in question. Cur-rently, this is achieved by helicopters when conventional means of transport is no longer applicable.However, Unmanned Aerial Vehicles (UAV) can be seen as a more sustainable choice for the future.Furthermore, the absence of human beings in the aircraft increases the range and flexibility of thetype of missions that the UAV can perform. This is of paramount importance when it comes tonatural disasters, as the situation can very easily be escalated depending on the surrounding of thesite. As was the case with the recent earthquake to hit Japan that triggered a sequence of eventsresulting in the meltdown of a nuclear plant.

This particular Design Synthesis Exercise (DSE) project deals with the design of an UnmannedAerial Vehicle (UAV) to provide humanitarian aid to the residents of areas struck by natural disas-ters. This report aims to provide a detailed overview of the results obtained from the entire designprocess of the DSE. The emphasis lies on the outcomes of the detailed design phase of the project.

The UAV is designed to be more cost efficient and environment friendly with respect to mannedhelicopters. The most vital requirements of the UAV include, the ability to perform a vertical takeoff and landing maneuver, a reduction in the Carbon Dioxide emission by 75% with respect to aconventional helicopter and the ability to fit 10 UAV units, support system and staff in the A400Mtransport aircraft. The designed UAV, Flamingo complies with all customer set requirements and isdesigned according to CS-23 regulations.

The structure of the report is as follows. Chapter 2 gives a brief overview of the requirement drivingthe design, the market analysis, the sustainable development strategy and the preliminary conceptsthat led up to the design of the Flamingo. Chapter 3 describes the aerodynamic, stability and perfor-mance characteristics of the UAV. A detailed analysis of propeller design is included here. In chapter4, structural analysis of wing box, canard, cargo pod, landing gear and booms are performed. Finiteelement modeling have been incorporated in this analysis. Chapter 5 explains the theory involvedin the computation of the numerical simulation of the VATOL maneuver. Chapter 6 describes thedesign of the UAV subsystems, whereas chapter 7 depicts the operations and logistics of the UAV.Chapters 8 and 9 gives the conclusions and the recommendations respectively.

1

2 1 Introduction

Chapter 2

Design Aspects

This chapter states the design aspects of this project and the basic UAV configuration. First, theproject objective is defined in section 2.1. Then the customer requirements are discussed in section2.2. The preliminary concepts are shown in section 2.3. Then, the winner of the trade-off is discussedin section 2.4. Hereafter the mission profile will be elaborated in section 2.5. The sustainibility isfocussed in section 2.6. Market analysis is performed in section 2.7.

2.1 Project objective

The Design Synthesis Exercise is the final project of the BSc Aerospace Engineering at the TU Delft.During the exercise, 10 students work full time for 11 weeks on an aerospace related design.

In the design project, it is demonstrated that the students have the basic knowledge and skillsnecessary to accomplish a successful ‘paper’ design of an aerospace system. By completing the project,the students demonstrate the following (Melkert, 2011).

1. They have technical competence; the ability to apply knowledge.

2. They have design competence: perform conceptual design of an aircraft, integrate life-cycle and sustain-ability issues in the design.

3. They have effective communications skills: plan, prepare, deliver and assess meetings, oral presentationsand written reports.

4. They show a professional attitude towards each other and their clients.

5. They are able to work in multi-disciplinary teams.

6. They understand contemporary & societal issues in their work.

7. They exhibit life long learning attitudes and abilities.

The final deliverables of the project include the final report, a final review for the tutor and coachesand a presentation at the design synthesis exercise symposium at the faculty of Aerospace Engineering.

The group-specific assignment for group 6 of DSE spring 2011 is as follows:

Design a cheap and sustainable UAS to supply aid several times a day, to difficult to access disasterareas, in 11 weeks time within a group of 10 students.

The UAV is designed based on customer and designer imposed requirements. Hence, the requirementsare analyzed in detail with the use of a requirement discovery tree and a requirement traceability table.This is followed by a more thorough functional analysis. The functions and the requirements of theUAV help to illustrate what the final product should be capable of and what the customer expects

3

4 2 Design Aspects

from the product. At this point, the conceptual design phase commences. During the conceptualdesign the group is split into three sub-groups, each generating an individual concept. Preliminaryperformance calculations are made and engineering drawings corresponding to the conceptual designsare created. For all concepts an initial risk analysis is performed. The risk analysis is used in thetrade-off, when the most feasible design is chosen. Next to the risk analysis, also a general marketanalysis, operations and logistics analysis are performed and a sustainable development strategy iscreated. A trade-off is performed, once the concepts have been developed in substantial detail suchthat quantifiable parameters, especially concerning mission performance, can be compared to choosethe concept that is developed in detail.

When the final concept of the UAV is known, a detailed design of this concept is performed.During this phase, emphasis is put on the design of the UAV systems, improved aircraft design,control and stability, mission simulation and operational return of the UAV.

2.2 Emergency relief UAV requirements

There are two types of requirements that concern the UAS design- Customer requirements as statedby the customer and designer imposed requirements. This section explains both theses kinds ofrequirements as well as top level requirements and children requirements which are a consequence oftop level requirements. The compliance matrix is put in Appendix C.

2.2.1 Customer requirements

This subsection explains the customer top level and children requirements. The UAV should be ableto carry a maximum of 250kg of payload mass and 1 m3 of payload volume.It should be able to landon an area of less than 10m by 10m. This implies that the span of the UAV should not exceed 14.14m.The UAV should also be designed such that it can be disassembled and 10 such units, support systemand staff can fit in an A400M transport aircraft which implies that 10 UAV units, support system andstaff should be able to fit in a total volume of 340 m3. In addition the UAV, when carrying maximumpayload, must be able to travel 200km to deliver the payload and another 200km to go back to theairport without the payload. Each UAV should conduct at least 4 sorties a day. This maximum rangeand mission duration imply that the cruise speed of the UAV should be minimum 32 m/s. Anothercustomer requirement specifies that the UAV should be able to take-off in an airport length of 1000m. Moreover the payload must not be accelerated to more than 2 g which implies that no rockets orboosters can be used for power systems. The UAV should be able to operate during a wind speedof more than 7 Beaufort and due to restricted area at the disaster site be able to land and takeoffvertically. The UAV should be able to land on a maximum slope of 10◦ and during night have thecapability to land autonomously, which implies that the sensor system of the UAV should be able tooperate in the absence of sunlight.

Another important customer requirement is regarding costs and emissions. The direct operationalcost of a UAV should not exceed AC500 per sortie while the unit production cost should not exceedAC250,000. The UAV should operate in an environment friendly manner and have 75% less CO2

emission compared to a manned helicopter. This implies that the maximum fuel mass per sortieshould not exceed 28.9kg and this fuel mass calculation can be found in section 2.6. Moreover theUAV should comply with present CS-23 airworthiness regulations and should be designed by 10 peoplewithin 11 weeks of time.

2.2.2 Designer imposed requirements

This subsection explains the designer imposed top level and children requirements. The UAV shouldbe able to operate in several kinds of environments because disasters are not region specific, shouldbe able to land on rough surface and be able fly at altitudes of about 6000 m because it should be

2 Design Aspects 5

able to reach the highest settlements in the world. The highest altitude for the UAV to fly at, withoutpayload, is the altitude required to fly over all parts of the Earth which is 9000m and the UAV shouldstay operational within the temperature range of -51◦ to 57.8◦. Moreover the UAV should be ableto operate in very humid environments without the risk of corrosion and should be able to operatein desert and arid regions. All these imply that the engine of the UAV should be operational at thestated altitude, ceiling and temperature requirements and also that the payload should be properlyinsulated in order to safely reach the disaster site. In addition the material used for UAV shouldbe able to withstand all possible environments. The design of the UAV should be such that it isproduced, operated and disposed in an environment friendly manner so the use of toxic materialsshould be minimized. The UAV should be able to deploy the payload automatically and reach thedisaster site which implies that a proper deployment mechanism is needed to deploy payload and tobring a UAV to disaster site it should have an accurate reporting status and needs to be equippedwith a proper guidance, navigation and communication system. It should have a sufficient powersupply to perform the mission and be able to provide visuals which implies it should have a highresolution remote sensing system. Besides the UAV should conduct the mission safely which impliesthat people and system should be safe.

All these requirements with their unique identifier are listed in appendix C. Moreover, the com-pliance traceability is included in this appendix as well.

2.3 Considered concepts

To give an insight in the initial phases of the design process, this section will give a quick introductionof the concepts considered before selecting one final design. Each of these concepts has been developedup until the class II weight estimation (class II weight estimation is a method to check the operationalempty weight which is calculated by the preliminary weight estimation method) and after that a trade-off has been made. The description of each of the concepts is as they were designed at the time ofthe trade-off, so there will be some discrepancy between the Tailsitter described at that time and thefinal design presented in this report. In this section, after a short description of each of the conceptsa table can be found containing the performance of the aircraft for a design mission of 200km, ashort mission of 150km and a longer mission with reduced payload of 300km. After the conceptdescriptions, the trade-off will be discussed and the results can again be found in a table containingall trade-off criteria and scoring. From this table it will become clear why the Tailsitter is the conceptthat has been further developed.

2.3.1 The helicopter concept

One of the concepts is the conventional helicopter. It has a main rotor which is attached to a huband a hinge. To counteract the reaction torque on the fuselage which is created by the main rotor atail rotor is attached to the boom. There are no horizontal or vertical stabilizers used. To control theattitude of the aircraft a swash plate is used which translates the reciprocating motion into a rotatingone.

The engine is mounted inside the fuselage and the fuel tank is located in the fuselage as well.Since the payload is carried externally the fuselage becomes smaller compared to one which carriesthe payload internally. The landing gear which is attached to the fuselage is a skid landing gear.Skids are used mainly because they weigh less than wheels and are cheaper to produce. There aretwo gearboxes used, one in the fuselage for the main rotor and the other one in the boom for the tailrotor. An artist impression of this concept can be found in figure 2.1. The mission performance ofthis concept can be found in table 2.1.

6 2 Design Aspects

Figure 2.1: Artist impression of the Helicopter concept

Mission distance (km) Time Taken (min) Fuel Consumed (kg) Cost (AC)

150 120 29.1 79

200 156 34.6 99

300 246 40.1 117

Table 2.1: Mission performance of the Helicopter configuration

2.3.2 The tiltrotor concept

Before developing a specific tilt rotor configuration, the choice was made between a quadcopter tiltrotor and a twin rotor tilt rotor. The quadcopter concept was discarded because it is less efficientin hovering, as two large rotors have a greater propulsive efficiency than four smaller ones. Also, anextra engine would be necessary to prevent very complicated gearing systems. The quadcopter doeshave a very good stability in comparison to the dual tilt rotor. As a compromise, a dual tilt rotorwith additional fan was chosen.

The payload is placed near the tiltrotor’s center of gravity. In that case, the shift in center ofgravity after delivering the payload is minimal. The foldable tail is mounted on the aft of the fuselage,after a small fan which will provide extra pitching stability during hovering. The engine is then placedin the nose, to shift the center of gravity back on the wing. An artist impression of this concept canbe found in figure 2.2. The mission performance of this concept can be found in table 2.2.

Mission distance (km) Time Taken (min) Fuel Consumed (kg) Cost (AC)

150 101 65.3 101

200 129 85.5 130

300 191 130 195

Table 2.2: Mission performance of the Tilt Rotor configuration

2.3.3 The tailsitter concept

The Tail Sitter concept aims to combine the mechanical simplicity and cruise performance of a fixedwing aircraft with the VTOL capabilities of a helicopter without the complexity of a Tilt Rotor. Theconfiguration of this concept comprises of a wing under which the payload is carried. The wing caries2 engines, one on each side. From the wing, two booms run forward which carry a lifting canard.

This concept will perform the mission by taking off horizontally from the airport. When landingat the delivery site the aircraft will pitch up steeply and slowly transfer the lifting load from the

2 Design Aspects 7

Figure 2.2: Artist impression of the Tilt Rotor concept

Figure 2.3: Artist impression of the Tailsitter concept

wings to the propeller. Once the aircraft is completely hovering it will slowly descent and finally landon its tail. After having delivered the payload, it will take off vertically again and transfer to cruise.When back at the airport it will land conventionally again. An artist impression of this concept canbe found in figure 2.3. The mission performance characteristics of this concept can be found in table2.3.

Mission distance (km) Time Taken (min) Fuel Consumed (kg) Cost (AC)

150 121 29.1 50

200 151 34.6 55

300 211 40.1 62

Table 2.3: Mission performance of the TailSitter Canard configuration

2.3.4 Trade-off

For the trade-off several aspects of the UAVs performance have been considered. Based on howimportant these aspects are considered to be, a weight has been assigned to it in terms of a percentageof the total score. The result of the trade-off can be found in table 2.4. For some of these aspectsno quantitative expression was available at the time of the trade-off (e.g. “cargo flexibility” and“reliability”). In such cases the concepts were compared to each other and the highest score wasassigned to what was agreed to be the ‘best’ configuration and lower scores were assigned accordinglyto the other two concepts. Note that the Tilt Rotor got a 0 score on manufacturing cost as it exceeded

8 2 Design Aspects

Criterion Weight% Helicopter Tiltrotor Tailsitter

200km Mission 24% 4.158 3.844 4.56

150km Mission 8% 3.575 3.17 4.47556

300km Mission 8% 2.82 1.652 3.026

Manufacturing costs 20% 5 0 4.48

Ability to bring back cargo 5% 3 5 3

Maximum altitude 5% 3 3.8 3.6

Reliability 5% 3 2 4

Risk 5% 3 2 4

Cargo flexibility 10% 5 5 4.25

Turnaround time 10% 3 3 5

Total weighted score - 3.610 2.448 4.246

Table 2.4: Trade off criteria and weights and scores for each concept

the maximum production cost of AC250, 000.

As can be seen, the Tailsitter concept wins by a significant margin. This is mostly due to the factthat it has the lowest operation cost and the fastest turn around time.

2.4 Features of the Flamingo

This section explains the general overview of the humanitarian relief UAV which has won the concep-tual trade-off. This concept is called the ‘Flamingo’ and will in the rest of the report also be referredto as such. First the general configuration is presented after which the VATOL capability and thecharacteristics of the cargo pod are explained. The details of the results of the efforts done by theproject team in the past few weeks will be presented in the following chapters.

2.4.1 Configuration

Figure 2.4: Artist impression of the Flamingo

An artist impression of the Flamingo can be seen in figure 2.4. As can be seen, the UAV is anairplane with a canard and two wing mounted propellers. The propellers are driven by one Wankelengine in front of the fuselage. Large vertical stabilizers provide lateral stability during flight. Thetotal span is about 7.5m, the length is about 5m and the height of the UAV without landing gear andpayload is approximately 2m. The airplane has a maximum take off weight of 720kg and a payloadcarrying capability of 250kg. The propellers are a compromise between helicopter propellers and

2 Design Aspects 9

conventional airplane propellers and have a diameter of 2.6m and are able to provide a maximumthrust of 7300N .

Detailed technical drawings can be found in appendix A and a project schedule for the remainingof the development of the UAV can be found in chapter 7.

2.4.2 VATOL capability

The tailsitter UAV has unique vertical attitude take off and landing, or VATOL, capabilities. VATOLcapable aircraft are usually referred to as tailsitters because they ’sit’ on their tail after a successfullanding is performed. The VATOL maneuver enables the UAV to land in areas without a runway,thus paving the way to humanitarian aid to nearly impossible to reach disaster areas. A VATOLis performed by making a pitch up maneuver. This means that the landing procedure starts with ahorizontal flight after which a canard deflection increases the pitch angle. The pitch angle is increaseduntil it has reached 90 degrees, the nose of the UAV now points upwards. Now the flight speed isreduced to zero and hovering flight like a helicopter starts. Attitude control is maintained by theelevons and rudders which feel an effective propeller induced air velocity. This part of the landingmaneuver requires that the thrust is equal to the weight of the UAV. When a stable hovering conditionis reached, thrust settings can be lowered and a vertical attitude descend can be initiated. When zeroaltitude is reached, a gentle touchdown enables the UAV to release its payload and take off again.The take off maneuver consists of applying maximum thrust, gaining sufficient altitude and flightspeed, and then easily pitching down in order to eventually perform conventional flight again.

2.4.3 Cargo pod

The cargo pod is a part which deserves special attention in this section. The cargo pod is able tocontain payload up to 1m3 and 250kg. Moreover, it is environmentally friendly and serves as a landinggear during the VATOL maneuver. The pod is made out of white spruce, which is cheap, strong andeasy to process. This is done because the cargo pod is left at the disaster site. After the touchdownof the VATOL maneuver, the pod with the payload is released and the UAV is able to take off witha significantly increased T/W ratio.

2.4.4 Safety

During the VATOL maneuver, the safety of the FLAMINGO and the surroundings of the landing sitemust be ensured since this is an uncontrolled environment. Especially because of the propeller heightof 1.5m during this vertical attitude maneuver, it must be ensured that there are no people on thelanding site. A LIDAR system is used to determine the landing site clearance. If the landing site isclear, the UAV makes sure that the contact with the ground is limited to the minimum time neededto unlock and release the payload. This can be a matter of seconds. This fast payload deploymentprocedure will enable the UAV to quickly hover to a safe transition altitude again.

2.4.5 Flexibility

The cargo pod presented in this report is primarily designed for emergency relief missions. For othermissions, other cargo pods can be designed according to requirements set by that particular mission.For example, for a long endurance observation mission, a cargo pod which contains an extra fueltank and observation sensors can be designed. This cargo pod can be made more expensive anddurable compared to the cargo pod for humanitarian relief missions since it is not disposed duringthe mission. Moreover, it can have a more aerodynamic shape, improving the flight characteristics.Another possibility is the use of the UAV as a light transport aircraft. The Flamingo then has adurable, aerodynamically shaped cargo pod which can be reused over and over again.

10 2 Design Aspects

2.5 Mission profile

In this section an overview of a typical mission is described. However, the UAS will be capable ofperforming various missions, which all might have a different profile. In this section ‘mission’ refersto an entire operation of the whole UAS, whereas ‘flight mission’ applies to a sortie of one UAV.The functional breakdown which describes the main characteristics of one general emergency reliefmission can be found in chapter 7.

Preparations

The main goal of the UAS is to supply aid to disaster stricken areas quickly. This implies that themission starts as soon as the help of the UAS is required after a disaster has happened. The UAS mustthen be transported to an airport nearby the disaster area quickly. To perform the transport fast,the total UAS including 10 UAVs, support system and personnel should fit in one A400M transportaircraft.

After the A400M has landed at a normal airport near the disaster site, the preparations forperforming the flight missions are initiated. The support system must be set up, the UAVs need to beassembled and the cargo pods can be loaded. Before and during the preparation phase, informationcan already be acquired about the cargo type and destinations.

2.5.1 Flight mission

As soon as all preparations are finished, the flight missions can be performed. A typical flight missionis bringing a payload of 250kg to a destination site 200km away from the airport. Before take-off theUAV systems should be checked to work correctly, after which the UAV can take-off conventionallyfrom the airport with a cargopod attached, and cruise to the destination.

On site operations

When the UAV approaches the landing destination, it uses its sensor system to create a virtual mapof the environment. A landing area of at least 10m by 10m is then chosen, where the landing willbe performed in a wind speed of up to 7 Beaufort. The VATOL maneuver and cargo pod releaseare performed as described in section 2.4. After the VATOL maneuver, the FLAMINGO switchesto conventional flight mode for the cruise back to the airport (without payload). The landing at theairport will be conventional again, after which the UAV can be prepared for another flight mission.The flight mission with respect to performance parameters is presented in section 3.7.

Additional mission characteristics

According to the requirements in section 2.2 this typical mission may not cost more than AC500 persortie. This cost comprises mainly of fuel, cargopod and communication data handling costs. Duringthe mission there should be 75% reduction in CO2-emission when compare to a manned helicopterperforming the same mission.

2.6 Sustainable development

A sustainable UAS is desirable for both ideological and practical reasons. Especially reducing theamount of fuel consumed for delivering emergency goods can be vital as fuel will be extra scarce in adisaster stricken area. The cargo pod, which is to be left behind on the landing site is another consid-eration. This section will discuss all the aspects considered to make this aircraft a more sustainablealternative as compared to conventional ways of transporting emergency goods.

2 Design Aspects 11

Mission phase Time required Fuel required

take off and landing 20 minutes 22.7l

cruise 111 minutes 126l

total 131 minutes 148l

75% CO2 reduction N.A. 37.1l

Flamingo 149 minutes 36.5l

Table 2.5: R-44 Mission profile

2.6.1 Manufacturing the UAV

When designing the aircraft, attention is paid to the materials chosen. As will be discussed in chapter4, most of the Flamingo is designed using metals. Only the booms carrying the canard and the cargopod are made of other materials. Primarily metals are used because metals can relatively easily berecycled at the end of life of the Flamingo1. The alternative would have been composite materials,however, until this moment, there are no real recycling possibilities for those materials.

2.6.2 Operating the UAV

The CO2 emissions should be reduced with respect to a conventional helicopter by 75%. This isgiven by requirement E.1. The Robinson R-44 is used as a benchmark. This helicopter has a similarpayload capacity to this UAV and is relatively cheap and will therefore form a good comparison. Thefuel consumption of R-44 is 68 liters per hour (Robinson Model R-44, 11-05-2011). It is assumed thata 75% reduction in fuel use will also result in a 75% reduction in emissions.

It is assumed that the 68 liters per hour is an average value for normal flight. The time the R-44requires to perform the cruise is found using the R-44s cruise speed of 216km/h (R44 Raven/ClipperSeries Helicopters, 17-05-2011). The time it will require this helicopter to take off and land at boththe airport and the delivery site is assumed to be 5 minutes for each phase, 20 minutes in total.The results for the fuel used for a ‘typical’ mission performed by an R-44 can be found in table 2.5.Flamingos performance is also given. The calculation of the mission performance will be elaboratedin section 3.7. As can be seen, the Flamingo meets this requirement with a small margin.

The cargo pod used to carry the payload is made out of white spruce. The structural reasoningwill be elaborated in section 4.3. The advantages for sustainability of the operation of the UAV isthat wood is naturally biodegradable or can be burned by the people on the disaster site with nodanger of toxic fumes.

2.7 Market analysis

The market analysis is performed in order to investigate the commercial potential of the UAV. First,an example of a natural disaster and its characteristics are explained as well as the potential cus-tomers. Then a market orientation is performed in which the capabilities of the UAS with respect tothese natural disasters are explained. Finally the positioning of the product with respect to possiblecompetitors is explained.

2.7.1 Disasters and customer

This section examines the practical application of the UAS after the catastrophic earthquake onJanuary 12, 2010 in Haiti. Since the earthquake caused several floods many areas became difficult toaccess by conventional means of transport. The U.S. military sent 23 Navy ships, 10 Coast Guard

1Regulations impose that recycled metals may not be used for aircraft anymore, however there are many otherindustries where these recycled metals can still be used

12 2 Design Aspects

ships, 264 fixed-wing aircraft and 57 helicopters to support humanitarian assistance and disaster relief.During the whole Haiti operation by the U.S. military, they supplied a total of 21 million kilogram ofhumanitarian aid consisting of water, food, radios and shelters Narrative History of Operation UnifiedResponse (2010). By assuming that every helicopter performed one operation a day, supplying aidweighing half the payload capacity of the helicopter, the helicopters supplied around 2 million kilogramof the total aid. When put into perspective of the UAV considered in this report, an A400M wouldbe able to land on three airports in Haiti located in Port-au-Prince, Jacmel and Cap-Haitien. Fromthese three airports every place in Haiti is accessible within a range of 200 kilometers which meansthat the UAV would be able to supply aid to every location in the disaster area.

Three main players can be distinguished when potential customers for a humanitarian relief UAVfor situations described above are investigated. These players are:

• International organizations

• Governmental institutions

• Non-governmental organizations and companies

There are several international institutions that could be interested in an emergency relief UAS.Institutes that quickly come to mind are for instance the UN, the NATO and the Red Cross. Theseinstitutes regularly send emergency relief aid to disaster areas. Organizations of this size often havea separate body that is specifically assigned to quickly react in case of disaster and this UAS canhelp them to react faster. Similar to the international institutes, individual governments mightbe interested in the UAS. Often emergency relief is coordinated through the defense ministry. Alsocountries with vast remote areas like the US, Canada, Russia and China might benefit from a UAS thatcan operate in remote areas, and archipelagos with small communities on different islands could havea use for this UAS. Finally, large NGOs and companies intervene in many different types of emergencyrelief missions. A UAS that costs about the same as one helicopter could be very attractive for theseorganizations. However this UAS is capable of more than just emergency relief. Oil companies forexample, might be attracted by the UAS’s capability of delivering supplies to their remote drillingsites. For similar reasons it might prove to be an effective transportation platform between remoteresearch centers for instance at the Antarctic Peninsula and the Arctic regions.

2.7.2 Humanitarian relief

In order to calculate what the UAS is capable of, the amount of beneficiaries in a specific area, thebeneficiary needs and the performance parameters of the UAS are taken into account (see section-sec:performanceoverview).

A beneficiary, a victim of a natural disaster, needs 0.75 kg of aid a day. This amount consists of0.6kg of food and 0.15kg of medicines and water purifying tablets. It is assumed that within a week,the stricken area will become accessible by other means of transport. So the UAV must be able toprovide humanitarian aid for a week, which gives a total of 5.25kg humanitarian aid per beneficiaryper week.

As one UAV is capable of providing 250 kg of humanitarian aid, within a range of 150km six timesa day, this means that one UAS consisting of 10 UAVs is able to provide 30, 000kg per day within anarea of 70, 000km2. To determine the number of beneficiaries which can be reached the populationdensity of the area is used.

Table 2.6 gives the number of beneficiaries which can be reached for a two day operation. Twodays is a typical time span which a human being is able to survive without (clean) water. Table 2.7shows the amount of beneficiaries which can be reached in a week. It is assumed that conventionalinfrastructure can be repaired within a week. In case the mission is extended for example for twoweeks, then accordingly the number of beneficiaries will double. It must be kept in mind that abovenumbers are applicable for average missions by one UAS consisting of 10 UAVs and operating from

2 Design Aspects 13

Population Density Possible beneficiaries Possible beneficiaries[people/km2] [% of total population]

10 5714 0.45%

50 5714 0.09%

100 5714 0.05%

Table 2.6: Number of beneficiaries for a two day operation

Population Density Possible beneficiaries Possible beneficiaries[people/km2] [% of total population]

10 20000 1.59%

50 20000 0.32%

100 20000 0.16%

Table 2.7: Number of beneficiaries for a seven day operation

one base only. For a disaster of a larger scale, for example the 2008 floods in Pakistan, more basesare needed to provide total coverage of the disaster area. It might be more convenient then to eitheruse more UAVs and/or spread the UAVs over multiple bases.

2.7.3 Positioning the product

For market analysis there are certain areas that need to be assessed, namely strengths, weaknesses,opportunities and threats. They determine the position of the product and represent all aspectsof the UAV design. Strengths and opportunities reflect the positive aspects of the UAV while theweaknesses and threats represent the downside. They are collectively termed as SWOT.The SWOTanalysis for the UAV is given in table 2.8 and is explained below.

Strengths Weaknesses

Ability to reach remote areas Can carry limited amount of payloadFlexible Can help limited number op peopleCheaper than alternatives No experience relief methodFaster than alternatives High risk design

Opportunities Threats

support helicopter missions Less trust in UAV developmentGrowing interest for other missions Competitions with reliable systems

Regulations

Table 2.8: The SWOT analysis of the Flamingo

Strengths

The strengths of the UAV include the ability of a UAV to reach remote areas that are inaccessibleby road and rail. Due to disasters people might be stuck in these regions and be in urgent needof survival supplies. UAV in this situation is a very practical solution. The UAV is also relativelycheap when compared to manned helicopters which are used for emergency relief purposes. This isa very important aspect as potential customers would want a system that costs less and fulfills themission’s objective. Moreover, the Flamingo is also able to fly faster than conventional helicopter andwill therefore reach disaster site more quickly. Flexibility is a very attractive aspect for the customerbecause it implies that the UAV is not restricted to perform only one kind of mission and can beadapted depending on the mission need.

14 2 Design Aspects

Weaknesses

A UAV can only carry a limited amount of payload. The size of a UAV is small which restricts theamount of payload that can be carried by it. On the other hand helicopters can be very large anddeliver a large quantity of payload which is very useful in regions that are densely populated. Since aUAV can carry less payload the number of people that can be helped is less. This is also a downsideof a UAV especially in situations when many people are in need of aid. Moreover, the design of aUAV is a novel process and designers don’t have any major prior experience. This increases the riskson UAV design.

Opportunities

A UAV can provide support to helicopter missions. In emergency situations it could be possible thatall available helicopters are not sufficient to provide aid, UAVs could then be used additionally tocomplete the mission. Or if the usage of helicopters is very costly they can be replaced by UAVsto make the mission more economical which increases the interest of potential buyers. In additionthe concept of UAV appeals to the customer because a UAV system can be used for various kindsof missions. For example a UAV could be used to supply aid but could also be used for observationmissions.

Threats

As UAV is a novel concept, many organizations may not trust it enough to readily implement it intheir mission. Besides, very reliable systems are available like helicopters thereby the customer ismore prone to the use of known and reliable alternatives than a new concept that has almost noexperience. Moreover, the UAV design has to be tested with respect to EU regulations. This is verytime consuming and till date there are not many regulations concerning UAVs.

Chapter 3

Aerodynamics, Stability andPerformance

This chapter discusses the investigation and optimization of the aerodynamic and static stabilitycharacteristics of the UAV as well as the performance characteristics. The combination of the canardconfiguration, the VATOL capabilities and the vast change in weight after releasing the payloadimpose several challenges.

Section 3.1 explains the methodology used for the aerodynamic analysis after which section 3.2discusses the aerodynamic characteristics of the UAV determined by this method. The static anddynamic stability for both conventional flight and hovering are discussed in sections 3.3 and 3.4respectively. The design of the propeller is elaborated in section 3.5. When all aerodynamic analysisis performed, the performance parameters and the mission profile characteristics can be determined.This is done in section 3.6 and 3.7. A final overview of all aerodynamic and performance parametersis given at the end of the chapter in section 3.8.

The outcomes of the aerodynamic and stability analysis are used for further structurally designingthe UAV and simulating the VATOL maneuver.

3.1 Aerodynamic analysis methodology

In order to analyze and optimize the UAV layout and flight characteristics the MATLAB basedprogram TORNADO (Melin, 2000) and the XFOIL based program XFLR5 (M. Drela, 2011)are used.While this section gives a general overview of the method and the aerodynamic model used, theprecise aerodynamic parameters used can be found in section 3.8.

3.1.1 The vortex lattice method

The vortex lattice method (VLM) is used to investigate and enhance the aerodynamic characteristicsof the UAV. The vortex lattice method divides the lifting surfaces of the aircraft in a number ofhorseshoe vortices creating a lattice of vortices used for lift and drag calculations. During the buildup of the aerodynamic model, it has been taken into account that the vortex lines are not coincidingwith a vortex lattice point. Otherwise it is not possible to mathematically analyze the aerodynamicsof the model. The use VLM implies several assumptions, namely:

• Homogeneous propeller induced velocity field

• The cargo pod has no influence on the flight derivatives

• Inviscid, incompressible and irrotational flow

15

16 3 Aerodynamics, Stability and Performance

• Thin lifting surfaces

The assumptions used have the implication that the method is only valid for small angles of attackand do not deal with stall characteristics. Please note that the flow is inviscid and therefore theparasite drag is underestimated by TORNADO.

3.1.2 Implementation in the simulation

The VLM aerodynamic model is used to determine the UAV flight coefficients which are input forthe simulation for a given state (e.g. airspeed, angle of attack). Because of the non conventionalmaneuvering done with the UAV and the accompanying (propeller induced) velocity field, the normalflight model cannot always be used. During hovering, the canard does not experience an effectiveincoming air flow, thus it will not influence the flight derivatives. To implement the absence of theinfluence of the canard in the simulation, an aerodynamic model without the canard will be used forthese flight regimes and a propeller induced velocity field is assumed. The rest of the model will beexactly the same as the model used for conventional flight. See chapter 5 for more details on theimplementation of the VLM model in the simulation.

3.1.3 The aerodynamic model of the Flamingo

First, the initial sizing done in the mid term design was used to build op the aerodynamic model ofthe UAV. After several iterations the final layout was determined as can be seen in figure 3.1. Belowthe different parts are described.

Figure 3.1: Aerodynamic model of the tailsitter using VLM in TORNADO

Main wing design

Since the fuselage of the tailsitter UAV is blended within the wing, the main wing is modeled as ablended wing body. To provide adequate space for the engine and the systems, the body has a NACA0025 airfoil and spans 1.04m. A ’transition’ phase of 24.4cm is used to merge the body with the mainwing. The main wing has a S4022 airfoil and flaperons spanning 2.2m and 25% of the local chordlength. The total span of the main wing is 7.45m and the wing tip has a twist of one degree.

3 Aerodynamics, Stability and Performance 17

Vertical stabilizer design

The size of the vertical stabilizers is, somewhat counterintuitive, not mainly determined by the aero-dynamic characteristics during the VATOL maneuver. Three conditions determine the shape of thevertical stabilizers. These are the location of the wing box of the main wing, the necessity to functionas the rear landing gear and the wish to put it as far after the center of gravity as possible to provideadequate yaw control during hovering. The first two conditions already lead to a surface area largeenough to provide sufficient yaw control during hovering. The airfoil used for the vertical stabilizersis a common used NACA 0018 airfoil. A 25% of the chord length is reserved for control surface.

Canard design

The canard size and location are optimized for stability and controllability. Just as on the main wing,a S4022 airfoil is used. The canard is not tapered and the span and chord of the canard are 2.6mand 0.4m respectively. The canard is located 30cm above the chord of the main wing to reduce theinterference during conventional flight.

Cargo pod design

The payload cargo pod is not directly modeled in TORNADO. Instead, a 3D model is made in XFLR5from which the drag coefficient is determined, see figure 3.2. Further details of the cargo pod can befound in section 4.3.

Figure 3.2: Aerodynamic model of the cargo pod using XFLR5

3.2 Analysis of the aerodynamics

This section discusses the outcomes of the aerodynamic model from the former section and concludeswith an overview of the main aerodynamic characteristics of the UAV.

3.2.1 CL − α and CL − CD curves

The CL−α curve of the UAV determined by TORNADO can be seen in figure 3.3. This figure showsthe curves for both the flapped and the unflapped wing. The main characteristics of the CL − αcurves can be found in table 3.1. Since TORNADO is limited to linear aerodynamic design the CLmax

and αstall are approximated using the DATCOM method (D.E. Ellison, 1965) while the exact stallcharacteristics are to be determined with more advanced modeling or experimental methods.

The CL − CD curve with and without payload can be seen in figure 3.4. The drag is estimatedusing two methods. One is the direct output of TORNADO, the other method is calculating the dragusing CDi = C2

L/πAe. As can be seen in the figure, the VLM analysis gives lower drag coefficientsfor higher values of CL. In order to not underestimate the drag, during the computation of theperformance parameters the drag coefficients from the general known equation for induced drag areused. The Oswald factor is assumed to be 0.9 for the main wing and 0.85 for the canard (Anderson,2005). The zero lift drag coefficients with and without payload are 0.0235 and 0.0092 respectively,

18 3 Aerodynamics, Stability and Performance

-­‐0.5

0

0.5

1

1.5

2

2.5

-­‐10 -­‐5 0 5 10 15

CL  

Alpha  [deg]  

CL-­‐alpha  curves  Re  =3.1e+6  

C_L-­‐alpha

C_L-­‐alpha,flapped

Figure 3.3: The CL − α curve of the tailsitterUAV

!"#$

"

"#$

%

%#$

&

" "#"& "#"' "#"( "#") "#% "#%& "#%' "#%(

! "#

!$#

!"#%#!$#&'()*#*+,!*+-./01.2345136.!789:;-8

*+,!*+-./01.2345136.!2<=>1=?3@A<.<BC3DE1@F

*+,!*+-./EDG.2345136.!789:;-8

*+,!*+-./EDG.2345136.!2<=>1=?3@A<.<BC3DE1@F

Figure 3.4: The CL−CD curve of the tailsitterUAV, with and without payload

Unflapped wing 20◦ Flapped wing

CLmax 1.60 1.94

CL,α=o 0.37 0.75

αL=0 −4.5◦ −9.0◦

αstall 15.9◦ 15.9◦

Table 3.1: Properties of the CL − α curve

this difference is calculated using the cargo pod model in XFLR5. (CL/CD)max with payload is equalto 14.8 and without payload is equal to 23.8. Please note that the zero lift drag coefficients are toolow, this can be explained by the fact that TORNADO assumes that the flow is inviscid as mentionedbefore. The ratio

(C3L/C

2D

)max

is 200 with payload and 321 without payload.

3.2.2 Main wing lift distribution

The main wing lift distribution can be found in figure 3.5. The curve shows clearly the influence ofthe fuselage (between −0.5m and 0.5m wing span) and the canards (at −2.5m and 2.5m wing span)on the main wing lift distribution. The lift distribution does not take the influence of the payload intoaccount. This would make the lift between −0.5m and 0.5m wingspan less. Integrating the curve forthe whole wing span gives the total lift of L = 6972N . Adding the positive canard lift gives a totallift of L = 8175.

3.3 Static stability

The static stability during flight is elaborated in this section. Unlike conventional airplanes, theUAV considered must also be statically stable during hovering flight. The release of the payloadon the landing site causes a great change in weight and a shift of the center of gravity which bothinfluences the stability behavior of the UAV. In order to gain a good insight in the static stability,the conventional flight and the hover flight stability is investigated both with and without payload.

3.3.1 The stability margin

The location of the center of gravity with and without payload and the aerodynamic center can befound in the technical drawings in appendix A. The locations of the center of gravity are determinedby the most up to date weight estimation. These include the weights of the different structural

3 Aerodynamics, Stability and Performance 19

0

100

200

300

400

500

600

-­‐4 -­‐3 -­‐2 -­‐1 0 1 2 3 4

Lift  [N]  

Span  location  [m]  

Lift  distribution  main  wing,  alpha=3deg,  V=55m/s,  Re  =  5.5e+6  

Lift  distributionmain  wing

Figure 3.5: The lift distribution of the main wing for α = 3◦ deg and V = 55m/s

parts, the engine, and all other UAV subsystems and their relative location. The payload causes thecenter of gravity to shift backward by 14.5cm, thus decreasing the stability margin. The location ofthe aerodynamic center is determined using TORNADO and shifts slightly with changing angle ofattack. The main wing is located as far back as possible in order to increase the stability margin.The center of gravity and the aerodynamic center were varied using different canard locations andsizes until an optimum was reached. The UAV is designed such that the stability margin is aroundzero with payload. A neutral or slightly unsteady equilibrium is allowed because the flight controller(see section 5.5.2) is able to cope with such a situation. The stability margin without payload allowsfor better stability and is xsm/c = 0.12.

3.3.2 Static stability during horizontal flight

Figure 3.6 shows the Cm−α curve both with and without payload. This figure shows two interestingfacts. First of all, for the UAV with payload Cmα is around zero, which indicates a neutrally stableflight. Secondly, around α = −3◦ the UAV shows a statically unstable behavior. This is due to theeffect of the canard on the main wing. At this angle of attack, the main wing is exactly in the wake ofthe canard. This unstable behavior at this angle of attack causes the UAV to go back to a stable angleof attack region again. Moreover, any unwanted instabilities can be canceled by the aforementionedflight controller by deflecting the elevons and canard.

3.3.3 Static stability during hover flight

To obtain static hovering stability, the free body diagram of the UAV during this phase is made, seefigure 3.7. It is assumed that there is no wind, the canard does not have a influence during hoveringand the propeller induced velocity is homogeneous, constant and only has a vector in the UAV x-direction. The forces acting around the center of gravity are the thrust force by the propellers andthe lift force by the main wing. The moment coefficient around the center of gravity by the mainwing is calculated for a range of elevon deflections using an aerodynamic model without a canard inTORNADO at a flight speed of 17m/s, the propeller induced airspeed. The moment coefficient dueto the thrust is taken constant with the thrust assumed equal to the weight.

The moment coefficient around the body m-axis as a function of elevon deflection is plotted infigure 3.8 with and without payload. This figure includes the moment coefficient induced by the thrustCm,thrust, which is equal to 0.06 and 0.35 with and without payload respectively. It was desired tohave the thrust vector in line with the center of gravity of the UAV with payload because the wingalready creates a moment around the center of gravity which is not negligible. As can be seen, there

20 3 Aerodynamics, Stability and Performance

-­‐0.15

-­‐0.1

-­‐0.05

0

0.05

0.1

0.15

-­‐10 -­‐5 0 5 10 15 20

Cm  [-­‐]  

alpha  [deg]  

Stick  fixed  static  stability  

Stick  fixed  staticstability  with  payloaddelta_can=-­‐2.7  deg

Stick  fixed  staticstability  w/o  payloaddelta_can  =  -­‐5.5  deg

Figure 3.6: The Cm − α curve of the tailsitter UAV with and withoutpayload

ThrustThrust

Lift

CG

Weight

Figure 3.7: Free bodydiagram ofthe UAVin hoveringphase

exists no moment around the center of gravity for δelevon = 10◦ (with payload) and δelevon = −10◦

(without payload). However, this elevon deflection causes the lift force acting on the body to havea Xearth-component. This component has to be counteracted by the thrust using a pitch angle ofθ 6= 90◦. Figure 3.9 shows the pitch angle needed to counteract the lift force in Xearth-direction withand without payload, which are 89◦ and 85◦ respectively. Of course, the propeller induced velocitywill change because of the inclination of the UAV (the thrust is not aligned with the weight any more,so T > W ), thus shifting the Cmδ,elevon-curve causing the elevon deflection to be slightly less.

The dynamic stability will be investigated by using the simulation (see chapter 5) since thrust isnot used as a variable in the aerodynamic model.

3.4 Dynamic stability

The dynamic stability during horizontal flight is investigated for one eigenmotion, the phugoid. Thisis only done for the phugoid to illustrate the aircraft behavior without controls. In reality, the controlsurface deflections will continuously be corrected by the flight controller.

The angle of attack and the velocity during the phugoid are plotted in figures 3.10 and 3.11 fora time span of 100 seconds. After 5 seconds trimmed flight a small negative canard deflection of0.03◦ is initiated. The plots show the unstable behavior of the UAV after that control input. Atapproximately 90s the maximum flight speed of 77m/s is almost reached while the maximum angleof attack during this eigenmotion is 3.2◦ and occurs at approximately 70 seconds.

3 Aerodynamics, Stability and Performance 21

-­‐0.6

-­‐0.4

-­‐0.2

0

0.2

0.4

0.6

0.8

-­‐40 -­‐20 0 20 40

C m  [-­‐]  

Elevator  deflection  [deg]  

Hovering  stick  fixed  stability  

Hovering  stick  fixedstability  with  payload

Hovering  stick  fixedstability  w/o  payload

Figure 3.8: The Cmδ,elevoncurve of the tailsit-

ter UAV with and without payloadduring hover

-­‐1500

-­‐1000

-­‐500

0

500

1000

1500

2000

80 85 90 95 100F X,earth  [N

]  

Pitch  angle  [deg]  

Hovering  pitch  angle  

Hovering  pitch  anglewith  payloadHovering  pitch  anglew/o  payload

Figure 3.9: Hovering pitch angle versus result-ing force in Xearth-direction withand without payload

Figure 3.10: Angle of attack of the Flamingoduring the phugoid motion

!"

!#

$"

$#

#"

##

%"

%#

&"

&#

'"

" (" $" %" '" )""

!"#$

%&'"

()$*"#&'"

+,-./)0"1*&2/3&*"

*+,-./01234.532

Figure 3.11: Velocity of the Flamingo duringthe phugoid motion

3.5 Propeller design

The proprotors play an important role for the tailsitter aircraft. They should provide thrust duringcruise flight like a propeller aircraft, but also provide thrust to hover during the VATOL maneuverlike a rotorcraft. The most important difference between these two phases is the velocity of the undis-turbed air relative to the aircraft: during hover this velocity is really low and maybe even negative,whereas during cruise flight the relative velocity can roughly vary between 40 and 60 m/s. The airfoilthat is selected is the VR7, which is an airfoil especially designed for VTOL aircraft.

There are quite some widely used theories involving the rotor hovering performance of a helicopter andthe propeller cruise performance of airplanes, but for the tailsitter the theories need to be combined.To do so, it is chosen to use the Blade Element Momentum (BEM) theory for propellers during thecruise phase and the BEM theory for rotors during the hovering phase. Also statistical data and theactuator disc theory are used for validation and verification. For the transitional phase, the requiredpower, delivered thrust and induced velocity are interpolated between the two phases.

The Blade Element Momentum theory discretizes the proprotor blades and calculates the local an-gle of attack at every section. Then, using the airfoil CL- and CD-curves the lift and drag contributionof all parts is calculated. Hereafter, the thrust and power are calculated.

Optimization is performed for shaft RPM, pitch angle, twist angle, taper ratio and solidity. Ahigher solidity turns out to improve the propeller performance significantly. From reference data of

22 3 Aerodynamics, Stability and Performance

Radius 1.3 [m] Twist hub 48 [◦] Taper ratio 1.0 [−]Solidity 0.11 [−] angle tip 0 [◦] No. of blades 3 [−]

Table 3.2: Basic proprotor configuration

PhaseRPM Pitch angle Required power Delivered thrust Induced velocity[rev/s] [◦] [kW ] [N ] [m/s]

VATOL 1400 10.0 75.6 3522 17.29

cruise 1000 23.5 20.23 222 0.455

VATOL wind 1400 10.0 78.3 3300 12.68

Table 3.3: Typical (single) proprotor settings and quantities

proprotor aircraft like tiltrotors a solidity of up to 0.11 seems feasible. When assuming a fixed gearratio between the engine and the proprotors some constraints are set between the cruise and VATOLRPM of the proprotors. Now by an iterative process, a twist angle and a taper ratio are found thatare able to satisfy the needs for both flight phases. For the CO2 requirement as stated in section 2.2(see requirement E.1), it is important that the required cruise power is kept low. For the VATOLmaneuver it is important that the required power does not exceed the available power as provided bythe engine in high performance setting and a Rich On Peak mixture.

After the iterative process, the basic proprotor configuration was determined as stated in table3.2. For the two typical flight phases, the corresponding proprotor settings are shown in table 3.3with their corresponding required power, delivered thrust and induced velocity.

For the wind speed of 7 Bft requirement as stated in section 2.2 (see requirement D.5), hoveringis also considered with a 17 m/s wind. The basic theory behind this is that the vehicle will haveto change its attitude to compensate for the drag: this changes the orientation of the thrust vectorand decreases the frontal area with respect to the wind. Using the previously described hoveringperformance calculations without wind, the configuration in windy weather is calculated. It turnsout that the thrust and power demands are only slightly higher during VATOL with wind. This ismainly because the canard is providing lift during hovering with wind.

The engine needs to be slightly oversized in consultation with the engine manufacturer, to fulfillthe power demands as stated in table 3.2. However, for these calculations, the in-ground effect hasnot been taken into account, which has a positive effect on the required power above the ground.

The simulation program as discussed in chapter 5 requires an input of airspeed and pitch angle,an output of required power, delivered thrust and induced velocity. To do so, air speed and pitchangle were varied and combined into three graphs. These graphs can be seen in figure 3.12. TheRPM setting is 1400 during VATOL and 1000 during cruise for this graph. The simulation programis now able to interpolate between these data points. Note that within the range of 0 to 15 m/s, noaccurate simulation data are available and this range is therefore purely interpolated. Note also thatonly relevant data points are shown: data points that are out of the range of interest are not shown.

The propeller efficiency varies during the flight. To give an indication: during cruise, the maximumpropeller efficiency that can be achieved is 0.88 at 50 m/s. At the typical cruise speed of 55 m/s,the propeller efficiency is 0.76. The propeller efficiency continues to decrease rapidly for higher cruisespeeds. During hover, it is common to express the efficiency using the Figure of Merit. For theFlamingo UAV, the FoM is 0.71.

3.6 Performance

The determination of the performance parameters follows from the aerodynamic and propeller anal-ysis. In section 3.6.1 the performance diagram is presented. The thrust to weight ratio diagram, of

3 Aerodynamics, Stability and Performance 23

Figure 3.12: Delivered thrust, required power and induced velocity of a single proprotor for varying pitchangles and velocity

prime importance during hovering, is presented in section 3.6.2. The rate of climb, stall speed, glidingrate of descent, flight envelope and take off field length are determined in subsections 3.6.4, 3.6.5,3.6.6 and 3.6.7. All performance parameters in this section are determined using performance theoryand equations from Ruijgrok (Ruijgrok, 1996).

3.6.1 Performance diagram

From the CL−CD-curve in section 3.2 and the total power available determination in section 3.5, theperformance diagram can be determined for different altitudes (Ruijgrok, 1996). The performancediagram, see figure 3.13, shows the power curves with and without payload. From these curves, thepower and velocity, which is the true airspeed (TAS), for maximum range and maximum enduranceare determined. Theoretically, the maximum flight speed can be determined from this diagram bydetermining the intersection point between the power available and power required. For the case offlying at 5000m altitude without payload, this would be over 150m/s. However, since the tailsitteris heavily overpowered for cruise flight the maximum flight speed is determined by the structurallimitations rather than the maximum power available, which is chosen to be 77m/s. Please note thatthe power available in reality decreases when the altitude increases but since the engine is overpoweredit will not affect the performance of this UAV.

3.6.2 Thrust to weight ratio

The amount of thrust decreases with altitude, therefore the maximum hovering altitude can be de-termined by plotting the maximum thrust to weight ratio versus the altitude without payload. TheT/W diagram can be found in figure 3.14. As can be seen, the maximum hovering altitude is 900mwith payload. This takes into account the fuel weight during VATOL of a typical mission describedin section 2.5. When calculating the maximum hovering altitude for a range of payloads up to thezero-payload situation, the maximum hovering altitude can be determined. The result of this analysis,which is done for maximum fuel weight, is in figure 3.15. Only altitudes up to 3000 are in the graph,since we would often only operate in this altitude range, and the results are less accurate when higheraltitudes are taken into account. Using extrapolation, this analysis gives an estimated maximumhovering altitude without payload of 8700m.

3.6.3 Range and endurance

The range R and the endurance E are given by equation 3.1 and 3.2, respectively in which thepropulsive efficiency ηj = 0.45 (see section 3.5) and the specific fuel consumption cP = 6.98E −

24 3 Aerodynamics, Stability and Performance

Figure 3.13: Performance diagram for the Flamingo

!"##

!"$

!"$%

!"$&

!"$'

!"$#

(

("!%

("!&

("!'

! )!! (!!! ()!! %!!! %)!!

!"#$%&

'()$

*+'(',-.$/01$

!"#$%&'()$2.%3,3$&+'(',-.$

*+,-./0123104-5/672/8

*+,9(

Figure 3.14: The thrust to weight ratio versusaltitude with payload

Figure 3.15: The maximum hovering altitudeversus payload weight

7N/Whr (given by the engine manufacturer). W1 and W2 are the begin and end weights of the UAVrespectively.

R =ηjcP

(CLCD

)max

ln

(W1

W2

)(3.1)

E =ηjcP

√√√√(C3L/C

2D

)max

2Sρ

[2√W2− 2√

W1

](3.2)

The maximum values for (CL/CD) and (C3L/C

2D) can be found in section 3.2. The range and endurance

at sea level can be found in table 3.4.

3 Aerodynamics, Stability and Performance 25

With payload Without payload

Endurance [hr] 4.1 5.2

Range [km] 589 944

Table 3.4: The range and endurance of the UAV

3.6.4 Rate of climb

The quasi steady rate of climb RCs is determined by dividing the excess power by the weight of theUAV. The excess power is calculated by the difference between Pa en Pr in figure 3.13. Thus equation3.3 is used for a range of flight speeds. The results, again with and without payload are plotted infigure 3.16. As can be seen, with payload the maximum rate of climb at sea level is 12m/s with aflight speed of 35m/s. For climbing with payload, the flight speed stays the same but the rate ofclimb almost doubles to 22.5m/s. With increasing altitude, the maximum rate of climb decreaseswhile the flight speed increases. Please note that the flight speed is the true airspeed (TAS).

RCs =Pa − PrW

(3.3)

Figure 3.16: The rate of climb versus altitudewith and without payload

0

5000

10000

15000

20000

25000

0 20 40 60 80 100 120

altitud

e  [m

]  

Flight  speed  [m/s]  

Stall  speed  versus  altitude  

Stall  speedwith  payload

Stall  speedw/o  payload

Figure 3.17: The stall speed versus altitudewith and without payload

3.6.5 Stall speed

The stall speeds for the UAV for different altitudes can be found in figure 3.17 and are determined byusing equation 3.4. The stall speeds take into account the maximum lift coefficient of the 20◦ flappedwing. The zero altitude stall speeds are 22m/s and 28m/s without and with payload, respectively.With increasing altitude, also the stall speed increases because of the decreasing air density.

Vstall =

√2W

ρSCL,max(3.4)

3.6.6 Gliding rate of descent

When an engine failure occurs, the gliding rate of descent RDgliding of the UAV and the flying altitudedetermines the distance which still can be flown to a safe landing area. The glide ratio is determinedusing equation 3.5 and assumes cosγd = 1 for the flight path angle during descend γd.

RDgliding =

√W

S

2

ρ

(C2D

C3L

)(3.5)

26 3 Aerodynamics, Stability and Performance

The rate of descent for minimum drag is obtained for maximum C3L/C

2D. This ratio with and without

payload is determined in section 3.2. The corresponding gliding ratios are RDgliding = 2.7m/s withpayload and RDgliding = 2.2m/s without payload.

3.6.7 Take off field length

Because of the huge amount of power available, the required runway length for take off is not de-termined by the maximum thrust during take off but the maximum moment the thrust is allowedto generate around the front landing gear. Assuming the most critical situation during the takeoff maneuver (without payload and at the beginning of the take off run) for the thrust T holdsT < Wxlg/zlg with xlg and zlg the distances of the main landing gear to the center of gravity (seeappendix A). Furthermore, with assuming a rolling friction coefficient of 0.05 for asphalt/concreterunways, the V/a-V diagram can be plotted from which the runway field length can be determined(see figure 3.18). Integrating the V/a-a diagram from zero to the lift off speed gives the total takeoff field length required. The lift off speed VLOF is given by equation 3.6 under the condition L = Wand W = MTOW , and is equal to 27.7m/s. The total required take off field length then becomes615m.

VLOF =

√2W

ρSCL,max(3.6)

Figure 3.18: V/a diagram during the take off run

3.7 Mission performance analysis

With all performance parameters known, the mission profile in section 2.5 can be investigated in moredetail. The goal of this section is to present the flight time, the fuel consumption and the total missioncosts. During the calculation of these characteristics, a discretized mission profile, a constant rate ofclimb and rate of descent and a fuel price of 2.74AC/liter (Rotterdam airport, 27-3-2011) are assumed.The mission profile consists of the take off and landing maneuver, climbing and descending, cruisingat 3000m altitude and the VATOL maneuver. The rate of climb and rate of descent are not equalto the maximum values determined in section 3.6.4, while the power setting during cruise is slightlylower than the setting for maximum range. This is both done in order to reduce fuel consumption ata cost of total flight time. The table 3.5 gives the flight time and fuel consumed during one 200km

3 Aerodynamics, Stability and Performance 27

range mission. The total costs, which include the fuel, the cargo pod and the communication datahandling are in table 3.6. A data transfer cost of AC2.94 per MB is used (BGAN Service Rate Pricingfrom Ground Control, 16-06-2011).

maneuver required time [min] fuel consumed [kg]

Take off and landing 6.0 1.1

climb and descend 30.0 2.4

cruise 108.5 20.8

VATOL 5.0 4.2

total 149.5 28.4

Table 3.5: The total time and fuel consumed for a 200km range, 3000m altitude mission

component costs [AC]

fuel 100

cargo pod 36

communication 15

total 151

Table 3.6: Costs per component of a 200km range, 3000m altitude mission

3.8 Overview of aerodynamic and performance parameters

In tables 3.7 and 3.8 an overview of all aerodynamic and performance parameters of the Flamingoare given.

28 3 Aerodynamics, Stability and Performance

Parameter value

Wing span b 7.45m

Wing area S 7.58m2

Taper T 0.56

Root chord cr 1m

Tip chord ct 0.5m

Mean aerodynamic chord MAC 1.42m

Wing span canard bcan 2.6

Wing area canard Scan 1.04m2

chord canard ccan 0.4m

Wing span vertical stabilizer bvert 2.04m

Wing area vertical stabilizer Svert 1.06m2

Root chord vertical stabilizer cr,vert 0.80m

Tip chord vertical stabilizer ct,vert 0.24m

Parasite drag CD0 0.009

Parasite drag coefficient with cargo pod CD0,cargo 0.0234

Maximum lift coefficient CL,max 1.60

Maximum lift coefficient, flapped CL,max,flapped 1.94

(CL/CD)max,withpayload 14.8

(CL/CD)max,withoutpayload 23.8(C3L/C

2D

)max,withpayload

200.1(C3L/C

2D

)max,withoutpayload

320.5

Stall angle of attack αstall 15.9◦

Table 3.7: Overview of the aerodynamic parameters of the UAV

Parameter value

MTOW 7140N

Empty weight 4150NTB -ratio 1.04/1.72

Wing loading 909/582N/m2

Maximum hovering altitude hhover 900/8700m

Maximum cruise altitude hcruise 17500/22000m

Range R 589/944km

Endurance E 4.1/5.2hr

Maximum rate of climb RC 12/22.5m/s

Climb speed Vclimb 35m/s

Stall speed Vstall 28/22m/s

Rate of descend during glide RDglide 2.7/2.2m/s

Take off field length stakeoff 615m

Table 3.8: Overview of the performance parameters of the Flamingo

Chapter 4

Structural Design & Analysis

In this chapter the structural design and analysis of the Flamingo will be discussed. First the designof the wing box will be explained in section 4.1. Followed by the structure of the canard in section4.2. Section 4.3 describes the design of the cargo pod and the release mechanism will be discussed insection 4.4. Then, the landing gear will be explained in section 4.5 followed by the description of thebooms between the wing and the canard in section 4.6. The control surfaces are explained in section4.7. Finally the shafts and gear boxes are described in section 4.8.

4.1 Wing box

In this section the design of wing box will be explained. The wing box is designed in such a way thatit is able to carry all the loads during a worst case scenario, which is assumed flying at a velocityof 77m/s at an angle of attack of 5 degrees. In this case the load factor is approximately 3 so thissituation can be compared to a turn with a 70 degrees bank angle.

4.1.1 Method used to determine normal stresses and shear flows

The first step is to determine all the forces that are acting on the wing. From the aerodynamic modelmade in TORNADO the lift and drag distribution on the wing were obtained for this situation. Thetotal lift and drag force are respectively 8540N and 853N from the wing tip to the root. The classII weight estimation gave an indication of the weight of every component on the wing. The thrustand torsion moment coming from the propeller were obtained using BEM theory. Figure 4.1 showshalf the wing plus fuselage and gives an indication of the forces that are acting on the wing box.The wing box is also drawn in this figure on which every boom has been assigned with a number. Inaddition, the figure shows where the cross sections are located and the reference frame that is used.Please note that forces are not drawn on the right scale and the reference frame selected is differentfrom the standard in figure 4.1.

After determining the loads, the wing box was sized to fit inside the wing. The wing box consistof a rear and front spar and a top and bottom plate with four spar caps in each corner. The frontspar is located at 0.1c and the rear spar is located at 0.4c where c is the local chord length. Oneshould note that at the fuselage the wing box is not located between 0.1c and 0.4c but from sectionC the dimensions increase with the same taper angle as between the wing tip and section C. Figure4.1 shows the location of the wing box. For calculating the normal stresses in the wing box idealizedboom theory (Megson, 2007) is used to convert the spars and plates into four booms. Since the wingbox is tapered the boom area increases from the wing tip to the root. Next, the wing is divided inmultiple cross sections which are shown in figure 4.1, and for each of these cross sections is checkedwhether the normal stresses are critical or not. This is done because since the wing box is tapered

29

30 4 Structural Design & Analysis

Figure 4.1: Forces acting on the wing box and their locations

the highest stresses are not necessarily found at the wing root. The normal stresses are caused bythe moment of the forces acting in the y-direction and by the moment of the forces acting in thex-direction. The normal stress (σz) in each boom at all the cross sections is found using equation 4.1.Here, M is the bending moment, I is the moment of inertia and x and y are the distances betweenthe boom and the local CoG of the cross section.

σz =

(MyIxx −MxIxyIxxIyy − I2

xy

)x+

(MxIyy −MyIxyIxxIyy − I2

xy

)y (4.1)

As expected the upper two booms experience a compressive stress and the lower two booms a tensilestress because the wing wants to bend upwards due to the high lift forces. Then the spar caps canbe sized in such a way that they do not buckle under these compressive forces. Using equation 4.2the critical buckling forces (Pcr) for the spar caps were found. Here, I is the moment of inertia,L indicates the length between the fasteners which connect the spar caps with the skin and E theyoung’s modulus which is 71GPa for the aluminium 7050 T76511 that is used. The reason why thismaterial is chosen is because it is commonly used for aircraft structural parts as plates, stringersand ribs. By iterating, the necessary amount of fasteners was found to prevent the spar caps frombuckling where also a safety factor of 1.5 is taken into account. This is present to guarantee that theaircraft can not only handle the limit loads but also the ultimate loads. Henceforth, the design willbe somewhat over sized to ensure safety.

Pcr =π2EI

(KL)2(4.2)

Thereafter, the torsion moment acting on the wing box was determined. The total torsion moment,which is the moment around the z axis, at each cross section is calculated by multiplying all the forcecomponents with their arm to the center of gravity of the wing box. Then by inserting the shear

4 Structural Design & Analysis 31

forces in the x and y direction the shear flows acting on each web were calculated using equation4.3. Here, S is the shear force, I the moment of inertia, td the web thickness, x and y the distancesbetween the boom and the local CoG of the cross section and Br the boom area.

qs = −(SxIxx − SyIxyIxxIyy − I2

xy

)(∫ s

0tdxds+

n∑r=1

Brxr

)−(SyIyy − SxIxyIxxIyy − I2

xy

)(∫ s

0tdyds+

n∑r=1

Bryr

)+ qs,0

(4.3)

4.1.2 Loads and stresses acting on each cross section

Table 4.1 gives an overview of all the loads, moments, stresses and shear flows that are acting on thewing box as each cross section. Please note that the safety factor of 1.5 is not taken into account. As

Cross section A1 A2 B C D E F

Cross section z-distance (m) 0.00 0.328 0.660 0.745 1.44 2.19 2.95

Chord length (m) 2.40 2.40 2.40 1.00 0.897 0.784 0.670

Thickness front spar (mm) 3.56 3.40 3.24 3.20 2.87 2.51 2.14

Thickness rear spar (mm) 3.56 3.40 3.24 3.20 2.87 2.51 2.14

Thickness top plate (mm) 3.34 3.12 3.04 3.00 2.69 2.35 2.01

Thickness bottom plate (mm) 2.78 2.66 2.53 2.5 2.24 1.96 1.68

Front spar length (mm) 66.7 63.8 60.8 60.0 53.8 47.1 40.2

Rear spar length (mm) 94.5 90.3 86.1 85.0 76.2 66.7 57.0

Top plate length (mm) 334 319 304 300 269 235 201

Bottom plate length (mm) 334 319 304 300 269 235 201

Shear force, Fx (kN) 0.427 0.389 0.350 0.340 0.260 0.174 0.0851

Shear force, Fy (kN) 8.92 9.52 8.62 6.38 4.69 2.95 1.30

Bending moment, Mx (kNm) −15.2 −12.6 −9.98 −8.78 −4.93 −2.07 −0.368

Bending moment, My (kNm) −0.745 −0.649 −0.539 −0.508 −0.232 0.130 0.0313

Torsion moment, Mz (kNm) 0.587 −2.77 −2.63 0.0150 0.0221 −0.245 −0.00182

Normal stress boom 1, σz (MPa) −82.6 −78.3 −71.5 −65.2 −51.0 −33.3 −9.57

Normal stress boom 2, σz (MPa) 104 98.5 90.0 82.3 64.0 39.3 11.2

Normal stress boom 3, σz (MPa) 138 131 119 109 85.1 54.9 15.8

Normal stress boom 4, σz (MPa) −126 −120 −109 −100 −77.9 −47.9 −13.6

Shear flow web 12, q12 (N/mm) 18.0 37.9 36.6 22.6 18.2 13.6 7.05

Shear flow web 23, q23 (N/mm) 219 262 250 182 149 108 55.5

Shear flow web 34, q34 (N/mm) 3.04 21.2 20.7 10.8 8.49 6.64 3.48

Shear flow web 41, q41 (N/mm) −231 −241 −228 −175 −144 −103 −53.1

Table 4.1: Loads, moments, stresses and shear flows acting on the wing box

expected the internal forces and moments increase as the cross section approaches to the root. Table4.1 also shows that the normal stresses in the two upper booms are compressive since they have anegative sign and the lower two booms are in tension. This is, as mentioned before, expected sincethe wing wants to bend upwards due to the high lift force. This also explains the high shear flows inweb 23 and web 41, because the top plate is completely loaded in compression and the bottom platecompletely in tension. The shear flows in web 12 and 34, which are the spars, are lower since thestresses change from compression in tension along these webs.

Also the number of fasteners required for the spar caps to prevent them from buckling werecalculated. Using equation 4.2, it is found that after each 22cm a fastener is needed on the spar capat boom 1 and after each 15cm at boom 4. However, the number of fasteners might increase becausethe spar cap should also be able to transfer the loads from the skin to the spars. Based on the resultscoming from those calculations also the number of fasteners required for the spar caps at boom 2 and

32 4 Structural Design & Analysis

a

b

ds

t

Figure 4.2: Overview of the variables used for the rib and stringer sizing

3 can be determined. This is however, because of time constrains, not done during this design phaseand is recommended to be done in the future.

4.1.3 Rib and stringer spacing

As discussed above wing box inside carries all loads acting on the wing. This wing box is a long,aluminum box, which is reinforced by stringers in the longitudinal direction, and ribs perpendicularto them. The length of the outer wing box considered, is 2.94m. The cross section decreases fromroughly 29.2 x 7.3cm to 20 x 4.8cm. To calculate the optimal amount of stringers and ribs, a MATLABprogram was written.

The input for the MATLAB program comes from analytical calculations performed in section 4.1.As explained in that section the stresses and forces are calculated at several locations on the wing.These locations are depicted in figure 4.1. Every section is assumed to have a constant width. Thewidth taken is the largest width of the section, to size it for the worst case, as larger plates are moreprone to buckling than smaller plates.

The MATLAB code is first used to compute the rib and stringer spacing on the top plate. Afterthat, a similar program computes the maximum stringer spacing for the bottom plate of the wingbox. The equations used come from (Bruhn, 1973) and (Megson, 2007).

Top plate

First, the program picks a thickness t, rib spacing a and stringer spacing ds. The maximum value fora is the segment length L, and the maximum value for ds is the distance between the front and rearspar b (see figure 4.2). For this combination of a, ds and t the critical buckling stress is calculated.This is done by first calculating the buckling coefficient k for both normal and shear stresses, whichis a function of a and ds. The critical buckling stress is then given in equation 4.4, where k can bereplaced by ks in order to get the critical shear buckling stress.

σcr =π2kE

12(1− ν2)

(t

b

)2

(4.4)

The ratio of the actual stresses and the critical stresses is then calculated, Rc for compression bucklingand Rs for shear buckling. To have an optimal sizing they have to meet Rc+Rs

2 = 1. A higher valuewould lead to buckling, a lower value leads to an oversized structure. The program then loops formany combinations of a, ds and t and saves all combinations for which yields that 0.99 ≤ Rc+Rs2 ≤ 1.From this list the optimum solution is selected by calculating the volume of material needed using anassumed rib and stringer thickness.

At this point it is calculated that the plate itself does not buckle under the loads, but it shouldbe checked that the stringers themselves do not buckle as well. It can be found that the force inthe stringers is distributed according to the ratio of the cross sectional area of the stringers and the

4 Structural Design & Analysis 33

Top plate Bottom plateSection t amax No. ribs dsmax No. stringers t dsmax No. strs

[mm] [mm] [-] [mm] [-] [mm] [mm] [-]C-D 1.9 100 6 70 4 1.6 100 2D-E 1.9 70 11 90 2 1.6 105 2E-F 1.9 110 7 120 1 1.6 120 1F-tip 1.9 390 3 160 1 1.6 140 1

Table 4.2: Rib Spacing

Figure 4.3: Top view of the rib and stringer distribution on the bottom plate

skin. Using the Euler buckling equation it is verified that the forces in the stringers do not exceedthe critical buckling force.

Bottom plate

For the bottom plate almost the same procedure is used. The biggest difference is that the rib spacingis predefined, because they were already determined for the top part. Furthermore the bottom platesare loaded in tension and shear, instead of compression and shear. However, they can still buckle dueto the shear loads. From (Bruhn, 1973) it follows that the equations for finding the optimum stringerspacing does not change, but Rc will now become negative due to the tensile loads.

Results

The final rib and stringer distribution is listed in table 4.2. According to the minimum rib and stringerspacing an optimal distribution was made manually, i.e. it was tried to make the stringers continuousand decrease the rib spacing to a minimum at heavy loaded parts. This is visualized in figures 4.3and 4.4.

4.1.4 Finite element analysis

The stresses in the wing structure mentioned thus far have been calculated utilizing idealized boomtheory. In an attempt to verify the aforementioned analytical results, various finite element modelswere built. Expertise in the field of finite element analysis was consulted to find the optimum softwareto be used prior to finite element modeling. Patran was the more accurate software choice, whichmakes use of MSC Nastran to analyze the finite element model (FEM). But due to its complexityof usage and due to the (relatively) low level of detail of the computed models during this project,CATIA’s Generative Structural Analysis workbench was chosen. For the goal of verifying analytical

Figure 4.4: Bottom view of the rib and stringer distribution on the top plate

34 4 Structural Design & Analysis

Figure 4.5: Overview of the outer wing box model

Figure 4.6: Cross sections of the wing box

results, this program is accurate enough. Please note that buckling analysis using finite element mod-els are beyond the scope of this project, so only deflections and stress distributions will be computed.

In order to perform finite element analysis in CATIA, an accurate three dimensional model of theinternal structure of the wing box was made. The analysis is performed for half the wing span be-cause of symmetry. Three finite element models are used to analyze the entire wing box. The firstFEM analyzes the outboard wing box, indicated as Section C to the wing tip in figure 4.1. Thetwo remaining finite element models analyze the inboard wing box given by Section A1 to SectionC in figure 4.1 for two different loading cases. One at the most extreme flight case and another atthe moment of impact during landing. Here, emphasis is put on making a structurally sound enginemount to minimize its displacement under loading.

Finite element analysis of the outboard wing structure

The three dimensional CATIA model of the outboard wing box was made in detail. It includes theassembly of all stringers and ribs in the tapered wing structure. As mentioned in section 4.1.3, theamount of ribs and stringers were calculated using a numerical simulation to ensure that the skinand the stringers do not buckle under the loads present. It was tried to make the model as realisticas possible, e.g. by making holes in ribs for stringers to pass through and the flanges for the ribs toconnect to the skin. The overall illustration of the wing box made in CATIA can be seen in figure4.5. Figure 4.6 illustrates a typical cross sectional view of the outboard wing box model. All the ribsand stringers were connected to the corresponding outer skins of the wing box using surface contactsin CATIA Assembly. When the three dimensional model is imported in to the Generative StructuralAnalysis workbench, these surface contacts are further defined as fastened connections. During theactual production process, various types of rivets and bolts will be used to connect the surfacestogether. Due to the already complex nature of the CATIA model and due to time constraints, it wasdecided not to include this type of connections in the model, but merely use fastened connections.This implies that inter rivet buckling properties are not considered in this analysis. However, testmodels were made to compare the stress distributions due to a fastened connection and due to rivet

4 Structural Design & Analysis 35

Figure 4.7: Stress distribution in the outer wing box using FEM analysis

connection. The overall stress distributions and deflections of both models were almost equal, exceptfor stress concentrations in the close proximity of the rivets. These results further reinforced thedecision to model rivets as fastened connections.

Note that due to issues with constraints present in CATIA, it was not possible to transfer theforces acting on the ribs directly to the spar webs, but rather the forces will be transferred throughthe top skin. This will not alter the results of the FEM significantly, as the model does not take into account the buckling of ribs.

All analytically computed loads are incorporated in the outboard wing box FEM. CATIA allowsthe application of a distributed load to a surface. As the upper skin was modeled as one surface,a total lift force of 7000N was distributed on the top plate instead of the detailed lift distributionillustrated in figure 4.1. As the force is distributed equally over the area, due to taper of wing theforce will increase on every segment closer towards the root. Because the applied force is not zeroat the wingtip in the model, the analysis is done for a worse case than in reality. All loads appliedto the FEM are listed in table 4.3. Please note that these loads derive from the analytical solutionexplained in Section 4.1.1.

The wing box at cross section C, figure 4.1, is modeled as a clamp and the aforementioned loadsare applied in the CATIA FEM at their corresponding lengths from the tip. The results can be seenin the figures below. Figure 4.7 illustrates the stress distribution along the length of the outboardwing box. Figure 4.8 illustrates the displacements of the wing box due to the applied loads. Theseresults are obtained after several iterations, where the mesh size has been refined to reduce the errorof the result (as estimated by CATIA). This error has been reduced to 12%. This error may seemhigh, but due to the complex nature of the model, it is very difficult to reduce this error further.Still, the maximum stresses are almost half the yield stress of the material, so the margin for plasticdeformation to occur is very high.

Type of force Force [N] Distance from tip [m]

Lift 7000 -

Propeller and fin weight 240 1.50

Wing weight 241 1.96

Drag 323 1.52

Thrust 700 1.50

Table 4.3: Loads applied to the FEM of the outboard wing box

Please note that the maximum stress of 248MPa occurs at the interaction point of the top skin andthe web of the spar. Such high loads are probably not an accurate indication of the true stresses in

36 4 Structural Design & Analysis

Figure 4.8: Displacement of the outer wing box using FEM analysis

the wing box, but originate from errors caused by discretizing the structure in a mesh by CATIA.The FEM also shows a high variation of stress through the height of the spar webs as shown by figure4.7. This is due to the aforementioned remark about the ribs not being directly connected to the sparwebs but being connected to the upper and lower skin of the wing box.

The stresses in both the top plate and the bottom plate are the highest in Section CD, as thebending moment is highest there. From the finite element analysis, the average stresses at SectionC range from 110MPa to 135MPa in the top and bottom plate respectively. The analytical resultsyield the maximum compressive stress at section C to be 100MPa in the booms of the top skin andthe maximum tensile stress to be 109MPa in the booms of the lower skin. The differences in theresults derive from the theory used in the computations. In the analytical approach stresses are onlycalculated in the booms, whereas the FEM analysis calculates the stresses at every location on thecross section of the wing box. However, the results are in the same order of magnitude, from which itcan be concluded that the idealized boom theory is applicable in this case. Analytical solution makesuse of idealized boom theory and in doing so approximates the area of the top skin and the bottomskin by two booms. Because this idealization does not take ribs into account, it needs a larger crosssectional area to carry the loads. The FEM model uses a more accurate model with a smaller crosssectional area, which causes higher stresses. Another source of errors between the models derivesfrom the fact that an average lift distribution was used in the FEM, where as the analytical modelutilized a more accurate lift distribution.

The maximum deflection is calculated to be 130mm at the wing tip. Please note that this is only thedeflection of the outboard wing section. The deflection of every part of the wing box can be found infigure 4.8.

Finite element analysis of the engine mount and internal wing structure

The wing box of the inboard wing section (Section A1 to C) must not only carry the loads listed intable 4.3, but also loads from the engine, canard and the landing gear. Note that only the weight ofthe engine is considered. The vibrational loads are assumed to be minimal, especially with enginebeing a Wankel engine. As mentioned before, a FEM analysis will be performed for two cases: oneat extreme flight conditions (maximum loads coming from the canard) and another during landing.The emphasis of these finite element models lies in ensuring that the engine mount is structurallysound and displaces very little due to the loads. The internal fuselage structure was not designed indetail, but a general overview can be seen in figure 4.9. Two booms are attached to the wing box atone side, and will support the canard booms at the other side (not visible). At the middle and thefront of the fuselage two cross bars connect the booms. They reinforce the structure by transferring

4 Structural Design & Analysis 37

Wingbox

Engine

Cross bars

Booms

Figure 4.9: General overview of the internal fuselage structure

Figure 4.10: Overview of the displacements of the inner wing box at the extreme flight conditions case

loads between the booms, and act as engine mount as well. Focus will be put on the booms mountedto the wing box, as they support the engine and should displace minimally. A large deflection of theengine from its original location could cause a failure of the gearbox or shafts.

The representation of the inner structure in the FEM model can be seen in figure 4.10. The di-mensions were derived from analytical calculations. The wing box starts at the root (Section A1) andis modeled to be clamped there. The wing box continues until the beginning of the outboard part(Section C). This is wing section that was excluded from the previous FEM. As can be seen by figure4.10, the finite model includes the connection of the boom (symmetrical analysis) to the front sparof the wing box and attachment points for the landing gear struts. The attachment of the engine ismodeled at one point (its CoG). Ribs have been added at the attachment points of the landing gearand the boom, to transfer loads from the front spar to the rear spar. In addition, the connectionpoint of the boom to the front spar is strengthened. Please note that in reality the booms must beproperly connected to the wing box, e.g. using a bracket. For the sake of the FEM analysis this wasnot necessary.

The FEM at the extreme flight condition phase models no landing gear force on the wing box,as the UAV is in the air. A shear force and a bending moment are modeled at the tip of the boom.These loads represent the most extreme lift loads created by the canard. In addition the weight ofthe engine is added at its center of gravity position. Since, this force acts downwards, it has the effectof relieving the moment at the root of the wing and at the boom connection to the front spar of thewing box.

The maximum stresses at the extreme flight condition are 227MPa and the maximum deflection47.9mm (illustrated in figure 4.10). This is within the strength limits of the material, however thedisplacement is rather large. In reality it will probably be smaller, as the displacement is mainly

38 4 Structural Design & Analysis

caused by bending of the booms. The mounting of the engine at several locations on the booms willhowever make the entire structure stiffer, which will decrease the total displacement.

Next, a FEM analysis is performed for the landing gear. In this case the forces acting on the landinggear at moment of impact are modeled in CATIA as calculated in section 4.5. The canard is assumedto generate equal forces as during cruise flight.

For this conditions, the maximum stress is 400MPa. This is very close to the yield strength ofthe used material. The maximum deflection in the landing case is calculated to be 22.8mm. Due totime constraints, further iterations were not possible. Thus, it can be concluded that the wing boxneeds to be reinforced in the center. This could be done by adding stringers to the top and bottomplate.

It is clear that the model of the inboard wing is in no means as accurate as the model developedfor the outboard wing box. This is due to time constraints. Some simplifications made, include themodeling of stringers as increased skin thicknesses in the top and bottom plates. This simplifiedmodel should be developed further, by computing the number of stringer and ribs required to ensurethat the top skin does not buckle and that other wing box failure modes are not met. Furthermore,the total engine load was assumed to act at the center of gravity location of the engine.

Henceforth, the results maybe of low accuracy. However, these results show that a feasible solutioncould be developed for the configuration of the UAV, without the need for an extremely heavystructure.

4.1.5 Material and weight

The whole wing box, including the ribs and stringers, is made of aluminum 7075 T76511. Aluminumis notable for being a lightweight metal and of the aluminum series the 7075 T76511 type has thehighest strength among aluminum alloys. Moreover, it is commonly used in aircraft structural parts.Table 4.4 shows the material properties for this type of aluminum. From the detailed CATIA drawing

Material Characteristics Values

Density 2810 kg/m3

Price 1.33AC/kg

Youngs modulus 71 GPa

Poissons ratio 0.33

Tensile strength 469MPa

Compressive strength 441MPa

Table 4.4: Material properties of aluminum 7075 T76511

of the wing box from cross section C to the wing tip, figure 4.1, a weight was found of 17.2kg. Theinternal part weighed 7.9kg. Based on this value an estimation of the total weight of the wing box of50kg was made.

4.2 Canard design

In this section the structural design of the canard will be explained in detail. The canard is designedfor the same worst case scenario as the wing. The span of the canard is 2.6m and the geometric chordlength is 0.4m. Since the canard is not tapered the geometric chord is the same through out the wholespan. The canard is connected to the booms which are connected to the wing by two hinges each ata location of 1/4 of the span of the canard (at 0.65m) from the tip. These hinges are connected tothe rear spar of the wing box of the canard. The forces acting on the canard are lift, drag and theweight of the canard. The lift and the drag distribution on the canard are obtained from the program

4 Structural Design & Analysis 39

Figure 4.11: The wing box cross-section of the canard

Figure 4.12: Moment diagram due to lift (left) and moment diagram due to drag (right)

TORNADO. Now the forces acting on the canard are known, the internal structure of the canard canbe designed. Inside the canard, a wing box is placed which has a front and back spar, top and bottomplates and spar flanges at each corner of the cross section to connect the spars with the plates. Thefront spar is placed at 0.1c and the rear spar is placed at 0.4c. In the wing box cross section of thecanard idealized boom theory is used to convert the area of spars, plates and the spar flanges intofour booms. So at each corner there is a boom placed and it is assumed that all direct stresses arecarried by the booms while the skin is effective only in shear. The cross sections of the wing box ofthe canard for real and the idealized case are given in figure 4.11.

Since the canard is not tapered the boom area remains the same through the whole section. Nowthe cross section of the wing box of the canard is known, the direct stresses can be calculated at eachboom at different sections in the span of the canard. The canard is cut at 1.3m and the calculationsare done only in the first 1.3m, since the other part is symmetric. The section 1.3m is further cutinto four section each at a length of 0.325m and normal stresses acting at the cross-section of each ofthese length are calculated using equation 4.1 which is given in section 4.1. The bending moment iscaused by the lift and the drag forces. As expected the bending moment is high right before the hingeand it remains the same right after the hinge. The moment diagrams due to lift and drag are givenin figure 4.12. In table 4.5 the normal stress at each boom, the boom area and the normal forces atdifferent section are given. The calculated normal stresses are multiplied by a safety factor of 1.5.

4.2.1 Shear stress

As mentioned before the skin carries the shear stress. The shear stress at different section should becalculated since the shear web can buckle due to higher shear stress than the critical shear stress.

40 4 Structural Design & Analysis

Section Mx My Boom σ 1.5∗σ Fnormal[m] [Nm] [Nm] [-] [MPa] [MPa] [N]

0.325 5.84 3.35

1 0.760 1.14 1442 -0.530 -0.800 -1023 0.840 1.26 1314 -1.10 -1.65 -173

0.65 171 13.41

1 20.0 30.1 3.82·103

2 -19.1 -28.6 -3.65·103

3 28.9 43.3 4.50·103

4 -29.8 -44.7 -4.67·103

0.975 38.8 3.35

1 4.56 6.84 8682 -4.3 -6.47 -8263 6.54 9.80 1.02·103

4 -6.77 -10.2 -1.06·103

Table 4.5: Bending moments, normal stress and normal force on booms at different section in the wingbox of the canard

The shear stress is calculated by calculating the shear flow first and dividing that by the thicknessof the skin. The shear flow is calculated using equation 4.3. The lift and the drag forces are actingat the aerodynamic center. To calculate the constant shear flow qs,0 the moment is taken at thecenter of gravity of the cross-section and the shear forces are placed at the center of gravity from theaerodynamic center. In this case there will be two torque acting at the center of gravity which arecreated by these two shear forces since these shear forces are moved from the aerodynamic center andduring the calculation of the shear flow, these two torques are also taken into account. The shearstresses are given in table 4.6. The shear stress is multiplied by a safety factor of 1.50.

Sect. Sy Sx tthickness q 1.5q τ[m] [N] [N] [mm] [N/mm] [N/mm] [MPa]

0.325 297 20.6

rearspar : 1.00 5.33 7.99 7.99bottomplate : 1.50 0.130 0.190 0.130frontspar : 1.00 4.41 6.62 6.62topplate : 1.50 0.390 0.580 0.390

0.650 588 41.3

rearspar : 1.00 10.3 15.4 15.4bottomplate : 1.50 0.520 0.780 0.520frontspar : 1.00 9.01 13.5 13.5topplate : 1.50 0.18 0.26 0.18

0.975 191 20.6

rearspar : 1.00 3.36 5.04 5.04bottomplate : 1.50 0.190 0.290 0.190frontspar : 1.00 2.93 4.40 4.40topplate : 1.50 0.0230 0.0340 0.0230

Table 4.6: Shear flow and the shear stresses

4.2.2 Stiffeners of the canard

Due to lift the top plate of the wing box will be in compression and the bottom part will be in tension.Due to compression the top plate of the wing box of the canard will buckle. To prevent bucklingthere are number of stiffeners placed. The number of stiffeners differ per section of the wing box sincethe normal load is not constant through out the whole canard. The normal force at each stiffeneris calculated and compared with the buckling force to check whether the stiffener will buckle. The

4 Structural Design & Analysis 41

Figure 4.13: The cross-section of the stiffener and the spar flange

buckling force is calculated using equation 4.2 and when the buckling force is higher than the normalforce acting on each stiffener, the stiffener will not buckle. The material that is chosen for the stiffeneris aluminum 7050-T76511 and the young’s modulus is equal to 71GPa (CES Edupack Software, 2011).The bottom plate is in tension and the number of stiffener that are needed on the bottom plate iscalculated using equation 4.5.

σ = PL/HA (4.5)

In equation 4.5 P is the tensional load, L is the length of the stiffener, H is the height of the wingbox which is equal to 0.031m (average height since the wing box cross section of the canard is notsymmetric), and A is the cross-sectional area of the stiffeners and the skin together. The tensionalstrength of aluminum 7050-T76511 is 469MPa (CES Edupack Software, 2011). Once the total areais known the skin area is subtracted from it which gives the total stiffener area. In figure 4.13 thecross-section of the stiffener is given. By dividing the total stiffener area by the cross-sectional area ofthe stiffener the number of stiffener can be found. The shear web can also buckle if the shear bucklingstress is less than the shear stress acting on it. This critical shear stress on the web can be foundusing equation 4.6 (van Baten and G.N.Saunders-Smits, January 2004)

τcr = 8.5E(t/d)2 (4.6)

In equation 4.6, t is the thickness of the web and d is the stiffener pitch on the shear web. In table4.7 the number of stiffeners at the top plate and the rear spar are given.In table 4.7 a is the number ofstiffeners in the top plate and the b is the number of stiffeners in the rear spar. The stiffeners in thebottom plate are calculated as follows. If the canard is cut at 0.325m, 0.65m and 0.975m, there areonly stiffeners needed on the bottom plate between the section 0.325m− 0.975m since the tensionalforces are high here because the hinge is placed in this section and from the calculation it is obtainedthat there are 2 stiffeners needed with a length of 0.65m. So in total there are 4 stiffeners of a lengthof 0.65m needed in the bottom plate of the canard. The cross-sectional area of this stiffener is givenin figure 4.13. Number of stiffeners in the front spar is the same as rear spar. Number of stiffeners

Sec. Fcomp Fcr,stif a Fstif,top τmax τcr b[m] [N] [N] [m] [N] [MPa] [MPa] [m]

-0.325-0.325 172 115 1(length = 0.650m) 15.6 7.99 12.5 3

0.325-0.650 4.67 · 103 459 2(length = 0.325m) 211 15.4 18.6 2

0.65-0.975 4.67 · 103 459 2(length = 0.325m) 211 15.4 18.6 2

0.975-1.30 1.06 · 103 459 1(length = 0.325m) 95.9 5.04 5.71 1

Table 4.7: Number of stiffeners needed on the top plate and the rear spar

42 4 Structural Design & Analysis

Figure 4.14: Cargo pod configuration

needed on the top plate is 10 with a length of 0.325m and 1 stiffener with a length of 0.65m, onthe bottom plate 4 stiffeners with a length of 0.65m, on the front spar 12 stiffeners with a length of0.031m and on the rear spar 12 stiffeners with a length of 0.031m. The number of ribs needed is 6.

4.2.3 Spar flanges of the canard

The spar flanges are placed at the corner of the wing box cross-section and it is important to checkwhether they buckle due to the normal force acting on it. Using equation 4.2 the buckling force onthe spar flange is calculated. In equation 4.2 the L is the length where a rivet is placed. To sparflanges not to buckle in section 0-0.325m a rivet is placed at 0.325m, in section 0.325m − 0.975m, arivet is placed at each 0.093m and in section 0.975m− 1.3m, a rivet is placed at each 0.1m.

4.2.4 Material and weight

The material that is chosen for the wing box, stiffeners and the spar flanges is aluminum 7050-T76511. This material has very high strength coupled with high resistance to ex-foliation corrosionand stress-corrosion cracking, high fracture toughness and fatigue resistance. The young’s modulusof this material is 71GPa, the density is equal to 2180 kg/m3 and it’s cost is AC 1.21 per kilogram.The top plate’s length, width and the thickness are 2600mm, 121mm and 1.50mm, respectively(figure 4.11) The bottom plate’s length, width and the thickness are 2600mm, 120mm and 1.50mm,respectively (figure 4.11). The width of the top plate and the bottom plate differ since the wing boxof the canard is not symmetric. The front spar’s length, width and the thickness are 2600mm, 24mmand 1mm, respectively and the rear spar’s length, width and the thickness are 2600mm, 36mm and1mm, respectively. The volume of the wing box of the canard is 0.0011 m3 and the weight is 3.1kg.The length of the stiffeners are given in the section 4.2.2 and the cross-sectional area is 18 mm2. Byincluding the stiffeners, spar flanges and the rivets weights, the total material weight of the wing boxof the canard is equal to 4.1kg.

4.3 Cargo pod

In this section the structural analysis of the cargo pod is explained in detail. The cargo pod needsto be designed such that it can transport payload safely and can handle stresses that may occur inthe worst case scenario that is under the load factor of 2. Moreover, the pod needs to be as light aspossible, easy to manufacture, cheap and as it is left behind at the disaster site it should not have anegative environmental impact and be biodegradable. The cross section of the pod and the respectivedimensions are illustrated in figure 4.14. Each plate has a thickness of 4mm. The highest loads occurduring landing. In a normal situation, the cargo pod lands on one of its surfaces which will be calledthe pivot in the rest of this chapter and two legs which can be considered to be the landing gears.The location of the legs can be seen in figure 4.14. The dimensions of the leg are given in table 4.8.The load is therefore evenly distributed during landing between the pivot and the legs. However inthe worst case scenario the pod might only land on its pivot or one of its legs. Therefore the pivotand the legs should be designed such that they can individually handle the loads.

4 Structural Design & Analysis 43

Parameter Value(mm)

Outer diameter 22

Inner diameter 10

Length 325

Table 4.8: Dimensions of the legs of the pod

4.3.1 Centre of gravity

One needs to ensure that the centre of gravity of the pod lies between the legs so that the pod doesnot tip over. This also needs to be ensured at a 10 ◦ slope. Calculations for the center of gravity showthat it lies within the acceptable range and does not tip over even at a slope of 10 ◦.

4.3.2 Stress analysis

The forces that act on the pod include the weight of the loaded pod which is 5128N and the dragforce due to side wind of 7 Beaufort which is 86.7N. The weight of the loaded pod is obtained bymultiplying the weight of the payload with a load factor of 2 for the worst case scenario and dragforce is computed by using equation 4.7 where CD is the drag coefficient of the pod and is 1.

D = CD1

2ρV 2S (4.7)

When the pod lands on one of its legs the normal stress is given by equation 4.8 while shear stress isgiven by equation 4.9.

σ =F

A(4.8)

τ =V Q

It(4.9)

The stresses computed are multiplied with a safety factor of 1.5 before they are checked for buckling.This ensures that the structure is able to handle not only the limit loads but also the ultimate loads.The buckling stress for the legs is computed by using (Bruhn, 1973) equation 4.10. The table 4.9 liststhe results of the computations.

σcr =π2KcE( tl )

2

12(1− υ2)(4.10)

Since the direct stress is below the critical stress the legs do not buckle. Moreover the shear stress in

Stress Value(MPa)

Normal stress 4.2

1.5 × normal stress 6.4

Shear stress 3.6

1.5 × shear stress 5.4

Critical stress due to compression 12.5

Table 4.9: Stresses on the leg of the pod

the leg is lower than the shear strength of the material and the tensile stress is lower than the tensilestrength of the material. Thus the structure does not fail in any way. On the other hand, if the podlands on pivot only, the stresses calculated are listed in table 4.10 . However since in the plate holesare present due to rivets stress concentrations occur and the increased stress at the hole boundary isgiven by (Calister, 2006) equation 4.11.

44 4 Structural Design & Analysis

σm = 2σ

√d

2( 1κ)

(4.11)

The critical stresses are again determined by using equation 4.10 where l now represents the lengthof the shorter dimension of the flat rectangular plate. The results are summarized in table 4.10. As

Stress Value(MPa)

Normal stress 1.2

1.5 × Normal stress 1.8

Shear sress 2.2

1.5 × Shear stress 3.2

Normal Stress around hole edges 2.5

1.5 × Stress around hole edges 3.7

Critical stress due to compression top plate 0.57

Critical stress due to compression bottom plate 0.98

Critical stress due to compression side plate 1.11

Table 4.10: Stresses on the pivot of the pod

the critical stress is below the stresses acting the plates will buckle. There are two possibilities toensure the material does not buckle either by using a material with a much higher Young’s modulusor by adding stiffeners. The number of stiffeners that are needed to prevent buckling are determinedby (van Baten and G.N.Saunders-Smits, January 2004) equation 4.12. σ is the highest stress actingon the structure multiplied with the safety factor.

n =lav√1.9Et2

σ

(4.12)

The table 4.11 lists the number of stiffeners for each plate of the pod. Thus the total number of

Plate Number of stiffeners

Top plate 8

Bottom plate 6

Side plate I 6

Side plate II 8

Table 4.11: Stiffeners needed for the plates

stiffeners needed per pod is 42. The stiffeners are L shaped with length 10mm, width 9mm andthickness 1mm.

4.3.3 Set up

The pod consists of thin rectangular plates which have inbuilt stiffeners as calculated above. Theplates can be folded along the fold lines as shown by figure 4.15. Thus the pod does not take a lotof space in the A400M. During assembly the adjacent plates are rotated to the required extent andriveted together by personnel. This can be done rather quickly and easily by means of a rivet gun.15 rivets on each side are used to connect two adjacent plates of the pod. The top plate is not rivetedbut is locked by means of a conventional locking mechanism. The pod in total weighs 13kg. Thus thepod can be set up very quickly and loaded with payload. To prevent the payload from moving thepod is filled with plastic air sheets or air bags. This restricts the center of gravity location in the pod.Due to riveting one needs to determine the stresses because of the rivets. The highest stresses are the

4 Structural Design & Analysis 45

Figure 4.15: Cargo pod fold lines

rivet/plate bearing stress which are determined by equation 4.13 and is determined to be 21.4MPa.Thus the material needs to be able to sustain the increased stresses due to riveting.

σbearing =Faxialtdn

(4.13)

4.3.4 Material selection

White Spruce Picea Gauca (longitudinal) is a type of wood that will be used for the pod. The reasonbeing it has a small density when compared to others, is biodegradable, can be landfilled and can becombusted for energy recovery. Spruce has been used for aerospace before and can be used withinthe temperature range of the mission. Moreover this type of wood is acceptable according to MIL-S-6073 military specification concerning aerospace structures made of wood (Military SpecificationMIL-S-6073, 06-11-1945). The properties of this wood are listed in table 4.12. The table 4.12 liststhe characteristics of the stated material.

Material Characteristics Values

Density 360 kg/m3

Price AC0.491/kg

Youngs modulus 9800 MPa

Poissons ratio 0.35

Tensile strength 5430 MPa

Compressive strength 3210 MPa

Minimum service temperature −73 ◦C

Maximum service temperature 140 ◦C

Table 4.12: White Spruce Picea Glauca longitudional properties

4.4 Cargo release mechanism

In this section the release mechanism of the cargo pod will be explained. For the stability of theFlamingo it is very important that the cargo pod is not allowed to move during the whole flight.Especially during the pitch-up maneuver the cargo pod should not move at all. Another thing thatshould be looked at is that for aerodynamic reasons the top plate of the cargo pod should be positionedagainst the fuselage. To meet these requirements a support system was designed where four L-shapedhook profiles stick slightly out the bottom skin of the fuselage. On the top plate of the cargo podfour rectangular shaped supports are placed which have the same dimensions as the bottom part of

46 4 Structural Design & Analysis

Figure 4.16: Release mechanism of the cargo pod

the hooks. Then, when the cargo pod is positioned at the right location underneath the fuselage thebottom part of the hooks will go through the rectangular shaped support. Next, all hooks will berotated 90◦ by an actuator. This is done in such a way that cargo pod can not move in any direction.Figure 4.16 shows the top plate of the cargo pod. Here, the bottoms of the hooks are already rotatedand they are indicated with the slightly curved black rectangles. These profiles are curved so itdecreases the possibility that the corners get stuck behind the rectangular supports on the top plate.Then, when the Flamingo has reached the target area and has landed on site the actuator will beused to rotate the hooks back to its original condition. After the hooks are shifted upwards inside thefuselage the cargo pod is no longer supported and the Flamingo can take off with leaving the cargopod behind. This movement of the hooks towards the fuselage will only by around a centimeter sincethe thickness of the hooks is in the order of millimeters. As can be seen in figure 4.16 the lower hookshave to be rotated before the upper ones because else there is a possibility that the cargo pod movesdown a little which has as a consequence that the lower hooks can not be shifted towards the fuselageanymore.To distribute the loads coming from this support system to the wing box four booms were designed.The two booms that are connected to the front spar have a outer and inner radius of respectively18 and 14 mm and the booms that are connected to the rear spar have an outer and inner radiusof respectively 9 and 8 mm. The difference between these booms are caused by the fact that thedistance between the upper two supports and the front spar is much larger compared to the lowertwo supports. In addition, the upper two supports needs to carry the full weight of the cargo podsince the lower to supports are released earlier as mentioned before.

4.5 Landing gear

The Flamingo has two different landing gears: main landing gear and the tail landing gear. Themain landing gear is attached to the wing and the tail landing gear is attached to the vertical fins.

4 Structural Design & Analysis 47

The landing gear is designed for a worst case scenario; during hard landing the aircraft lands on onewheel with a landing gear load factor of 2 and a side force. To design the shock absorbers the landingimpact is taken into account, this is explained in section 4.5.1. The main landing gear is placed undera certain angle and the wheel is placed about 200mm in front of the center of gravity position toprevent the aircraft from tipping over.

4.5.1 Shock absorber

The basic considerations that affect the shock absorber are the sink speed, the load factor on landinggear, the stroke and the shock absorbers. According to CS-23 regulations, the small aircraft should beable to withstand the shock of landing at a sink speed of 3.0m/s at design landing weight and 2.1m/sat maximum gross weight (EAS, 2003). Since in this case, the landing gear is designed for worst casescenario the aircraft lands with the maximum gross weight, so the sink speed of 2.1m/s (7 ft/s) isused. The landing gear load factor for a small utility aircraft is 2-3 (Currey, 1988). In this case,the landing gear is sized for a load factor of 2. The shock absorber that is used in the main and thetail landing gears are rubber shock absorbers. The reason why this shock absorber is used while theOleo-Pneumatic shock absorbers have highest efficiency and the best energy dissipation is because ofthe simplicity, low cost and minimizing the precision machining. Besides, this shock absorber is usedfor light weight aircrafts, which is the case here. The rubber shock absorber’s efficiency is 60%. Therubber is used in the form of disks and the thickness of each disk is limited by the thickness that canbe vulcanized to the plates or washers used to seperate the disks in the stack. So the thickness shouldnot be more than 1.5 inch (0.038m) (Currey, 1988). By using a central tube all the rubber disks arekept in line. The hole in the center of each disk is lined with fabric. During compression, these holesbecome smaller and the fabric contacts the tube and thereby absorbing some of the energy by way offriction.The stroke is roughly a linear function of the load factor and is the vertical distance movedby the wheels. The stroke distance calculation of the main landing gear is given in section 4.5.2.

4.5.2 Main landing gear

In this section first the shock absorber is sized and then the main landing gear is sized. The strokedistance and the strut length of the shock absorber are sized for the landing impact. The strokedistance can be calculated based on the fundamental work/energy relationship: change in kineticenergy is equal to work done. Applying that to a landing gear, a change in kinetic energy is equal toreduction of vertical velocity from sink speed to zero. In equation 4.14 the change in kinetic energy isgiven in the left side of the equation and the work done by different components is given in the rightside of the equation.

0− WVs2

(2g)= −S · nS ·NW − St · nt ·NW +W (S + St)− L(S + St) (4.14)

In equation 4.14 the first term on the right side of the equation is the work done by the strut, thesecond term is the work done by the tire, third term is the work done by the gravity and the fourthterm is the work done by wing lift. According to CS-23 regulations the lift created throughout thelanding impact is two third of the gross weight (EAS, 2003). In equation 4.14 the term S is the stroke(vertical wheel travel) in [m], ns is the shock absorber efficiency which is equal to 0.60 (Currey, 1988),N is the load factor which is equal to 2, W is the gross weight [N], St is the tire deflection and that isequal to 0.33 feet (0.10m) for small tires (Currey, 1988),nt is the efficiency of the tire which is equalto 0.47 (Currey, 1988), L is the lift [N ], Vs is the sink speed which is equal to 2.1[m/s] and the g isthe gravitational acceleration [m/s2]. When these values are substituted in equation 4.14 the strokedistance becomes equal to 7.78 inch (0.20m) and adding a 1 inch to the stroke for the inaccuracies thestroke distance becomes equal to 0.22m. The strut length of the rubber shock absorber is 2.5 timesthe stroke distance and that is equal to 0.56m (Currey, 1988).

48 4 Structural Design & Analysis

Parameters Unit At E in EG At T in AE At S in AE Strut DB Strut HC

Fnormal [kN ] 5.40 11.0 19.9 24.2 10.8

1.5Fnormal [kN ] 6.15 16.5 29.9 36.3 16.2

Lcuts [m] 0.1 1[m] 0.74 0.79[m] 0.93

Fcr [kN ] 5.85·103 388 1.82·103 50.2 26.6

σA [MPa] -318 -151 415 74.1 -41.4

σB [MPa] 288 119 -442 74.1 -41.4

σC [MPa] 228 400 84.6 74.1 -41.4

σD [MPa] -257 -433 -112 74.1 -41.4

τA-due Fz [MPa] 68.2 5.53 26.2 0 0

τA-due Fy [MPa] 34.3 14.2 3.78 0 0

τC-due Fz [MPa] 42.9 3.67 16.5 0 0

τC-due Fy [MPa] 54.6 21.4 6.01 0 0

τC-due torsion [MPa] 0 4.5 0 0 0

R1,inner [m] 0.0150 0.0300 0.0300 0.0100 0.0100

R2,outer [m] 0.0200 0.0350 0.0400 0.0160 0.0150

Table 4.13: The normal stress and the shear stress at different location in the landing gear

In figure 4.17 the main landing gear is sketched. The main landing gear is composed of a beam,two side struts, 2 hinges and a roller support. The beam and the struts are circular tubes. The beamis connected to the wing by a hinge, the front strut is connected to the wing by a hinge and backstrut is connected to the wing by a roller support. These two side struts are two force members.The front strut prevent the landing gear to move due to side forces and the back strut prevents thelanding gear to move due to normal force and the ground drag. When the landing gear retracts thefront sidestrut rotates with the main beam and remains at the same place. The back strut worksas a slider; during retraction it slides to the front and pushes the beam to go up (works as an actuator).

Since the landing gear is sized for the hard landing case, the landing gear load factor is equal to2; this means that the weight acting on the landing gear is twice the aircraft weight. As mentionedbefore the lift is 2/3 of the maximum gross weight (EAS, 2003). The normal ground force acting onthe wheel is equal to 8.68kN (Ng = W − L). According to CS-23 certification, tire sliding coefficientof friction should be 0.8 so the ground drag force is equal to 6.95kN (Dg = µr · Ng). CS-23 is alsostating that this drag force should not be less than 25% of the vertical force on the wheel (EAS,2003). In our case the ground drag is 80% of the vertical force. The side force should be equal to 0.83times the maximum gross weight (EAS, 2003) and that is equal to 5.40kN. These forces create normalforces, bending moments, torsional moments and shear forces in the structure which creates, directstress and shear stresses. It should be checked how high these stresses are and whether the landinggear material will be able to carry them without failure. First the unknown forces are calculated bytaking moment and forces equilibrium at the hinge point. Once the forces are known, the structurecan be cut at different location and the stresses can be calculated. The normal stresses are high atthe edges of the cross-section, so at each cross section the normal stresses are calculated at 4 points(two vertical and two horizontal).The normal stress is calculated using equation 4.15 (R.C.Hibbeler,2005). At each of these points the shear stress is also calculated and the equation used to calculatethe shear stress due to shear force is given in equation 4.9 in section 4.3 and the equation used tocalculate the shear stress due to torsion is given in equation 4.16 (R.C.Hibbeler, 2005).

In equation 4.15 N is the normal force [N], A is the cross-sectional area [m2], M is the bendingmoment [Nm], I is the moment of inertia [m4] and the y is the distance from the center of gravity tothe edge [m].

4 Structural Design & Analysis 49

Figure 4.17: Sketch of the main landing gear and cross-section of the beam

σ =F

A± My

I(4.15)

τ =Tr

J(4.16)

In equation 4.16 T is the torsion, r is the distance from center to stressed surface in the given positionand J is the polar moment of inertia of a hollow section. The points in each cross-section where thestresses are calculated are given in figure 4.17. In table 4.13 the normal stress and the shear stress atdifferent location are given. In table 4.13, the minus sign represents the compression. The materialthat is chosen for the landing gear is aluminum 7075-T73. The reason why this material is chosenis because this material has high strength, virtually immune to stress corrosion and most often usedfor the landing gear structures. This material has a young’s modulus of 72.5GPa, yield strength of446MPa, tensile strength of 510MPa and a compressive strength of 448MPa. The density is equal to2.81·103kg/m3. By calculating the volume in each section of the main landing gear the total volumeis found and that is equal to 0.00325m3. The weight is equal to 9.14kg. By giving a weight of 4kg forthe tire the total weight of one main landing gear is 13.1kg and the weight of the two landing gear isequal to 26.2kg.

4.5.3 Tail landing gear

The tail landing gear is attached at the vertical fin. The airfoil thickness of the bottom part of thevertical fin is increased slightly above the tire width and the wheel is placed inside the vertical fin.This wheel is attached to a beam which is connected to the rear spar of the wing box of the mainwing by a hinge in such a way the wheel moves with the vertical fin when the vertical fin is deflected.The beam has a length of 1.022m. The tail landing gear is sized for the maximum static load actingon it. To calculate the maximum static load, figure 4.18 is used. The maximum static load on thetail wheel is calculated as shown in equation 4.17 (Currey, 1988).

Nstatic,maxstrut =W (F − L)

2F(4.17)

In this equation W is the maximum gross weight. The calculated maximum static load in one tailgear is equal to 624N. The forces acting on the tail landing gear is the normal ground force, grounddrag and the side forces. According to CS-23.499 certification the normal ground force acting on

50 4 Structural Design & Analysis

Figure 4.18: The tail gear location (left) and the cross-section of the beam (right)

Parameters Values Unit Parameters Values UnitFnormal 1.40 [kN ] MXdrag -1.15·103 [Nm]

1.5Fnormal 2.11 [kN ] MYsideforce 1.00·103 [Nm]

Fcritical 41.5 [kN ] τA-due Fy 8.70 [MPa]σA -510 [MPa] τA-due Fx 4.70 [MPa]σB 503 [MPa] τC-due Fy 5.30 [MPa]σC 440 [MPa] τC-due Fx 7.60 [MPa]σD -447 [MPa] R1,inner 0.0100 [m]

R2,outer 0.0170 [m]

Table 4.14: The stresses in the tail landing gear

the tail wheel should be 2.25 times the static load which is equal to 1.40kN, the side force shouldbe 0.7 times the normal ground force which is equal to 983N and the drag force should be equal to0.8 times the normal ground force which is equal to 1.12kN. These forces are transfered through thebeam and the hinge to the wing box. The forces that are acting on the beam create stresses andthese stresses should be calculated to check whether the landing gear can handle these forces withoutresulting in structure failure. The beam is cut at 1.021m from where the hinge is placed to find thehighest moment and the stresses are calculated in this cross section at four different points. These 4points at the cross section are given in figure 4.18. The results obtained from the stress calculationare given in table 4.14.In table 4.14 the minus sign represents the compressive stresses. The material that is used for thelanding gear is aluminum 7075-T6. This material has a high strength and is often used for aircraftstructures (CES Edupack Software, 2011). This material has a young’s modulus of 76GPa, yieldstrength of 530MPa, tensile strength of 580MPa and a compressive strength of 530MPa. Thedensity is equal to 2.83·103kg/m3. The weight of one strut is 1.70kg. By giving a weight of 2kg forthe tire the total weight of one tail gear is 3.70kg and the weight of the two tail landing gear is equalto 7.40kg. The total weight of the landing gear (main +tail) is 33.7kg.

4 Structural Design & Analysis 51

4.6 Booms

In this section the structural analysis of the booms will be explained in detail. Two booms run forwardfrom the aircraft body to the canard. They are detachable from both ends. They are clamped at theaircraft body and are connected to the canard by means of hinges giving canard freedom to rotate.Therefore, they are designed such that they can sustain the loads from the canard. Since the canardis 0.3m above the aircraft wing, it implies that the booms run not only forwards but also upwardscarrying the lifting canard. The length of the booms is 2.3m each. The outer and inner diameters ofthe boom are decided by initially giving random values. After this stress distribution and deflectionof the booms is determined. Once these computations are done, an iteration is performed by changingthe inner and outer diameters of the boom such that the stresses at all cross sections remain belowthe critical stresses and also to limit the deflection of the booms to about 1 ◦. The procedure carriedout is as follows. First the deflection and stresses in the booms are determined under the forces innormal flight. Then the booms are checked if they do not fail in the worst case scenario. This isdone by finding the stress distribution under worst case scenario forces and comparing them with theyield strength of the material. To prevent booms from failing the stresses at the cross sections needto be below the yield strength of the material. The iteration of boom diameters is done to preventoversizing the structure. The inner and the outer diameters should just be sufficient to prevent thebooms failure and to restrict the deflection. The table 4.15 lists the inner and outer diameters of theboom.

Location Value [mm]

Outer diameter 70

Inner diameter 40

Table 4.15: Boom diameters

4.6.1 Stress analysis

This sub section explains stress calculation at different cross sections of the booms under normalconditions. The forces from the canard on the booms are 24N along the boom axis and 430N normalto boom axis. The side force due to 7 Beaufort wind is not taken into account for calculations becausethe force is very small.

The force normal to boom axis creates a moment at different cross sections. Therefore the stressat each cross section is the sum of the normal stress and the bending stress due to moment created,which can be determined by using equation 4.15 in section 4.5. A positive result from equation 4.15represents tension while the negative represents compression. One knows due to forces the upper partof the booms will be in compression while the lower part will be in tension. This can be verified bythe results which states negative stresses for the top of the booms and positive stresses for the bottomof the booms. The stress is determined at the top and bottom surface of the boom. In the table 4.16the forces and stresses at cross sections are listed. The calculated stresses are multiplied by a safetyfactor of 1.5. The compression stress due to force parallel to boom is much smaller than the criticalstress. Thus the booms do not break or fail.

The stress distribution is illustrated in figure 4.19 for both the top and bottom surface.

The stress distribution increases in the worst case scenario. The forces from the canard in theworst case scenario are 1353N perpendicular to boom axis and 97N parallel to boom axis. Equation4.15 is used again to determine the stress distribution at different cross sections of the booms. Thetable 4.17 illustrates the results obtained for the worst case scenario. From the results obtained onecan see that the stresses at each cross section remain below the yield strength of the material andthereby the structure does not fail.

52 4 Structural Design & Analysis

Z M IXX σtop σbot 1.5 · σtop 1.5 · σbot σcr[m] [Nm] [mm4] [MPa] [MPa] [MPa] [MPa] [MPa]

2.307 993 1052924 -33 33 -50 50 52

1.73 745 1052924 -25 25 -37 37 92

1.153 497 1052924 -17 17 -25 25 207

0.576 248 1052924 -8 8 -12 12 832

Table 4.16: Stresses on the booms

Figure 4.19: Stress Distribution

Z M IXX σtop σbot 1.5× σtop 1.5× σbot σyield[m] [Nm] [mm4] [MPa] [MPa] [MPa] [MPa] [MPa]

2.307 993 1052924 -104 104 -156 156 221

1.73 745 1052924 -78 78 -117 117 221

1.153 497 1052924 -52 52 -78 78 221

0.576 248 1052924 -26 26 -39 39 221

Table 4.17: Stresses on the booms worst case

4 Structural Design & Analysis 53

4.6.2 Material selection

This section explains the detail of a suitable material selection for the booms. In the class II weightestimation of the mid term design, the weight of the booms were approximated to be 5 kg each.The suitable material for booms should therefore have a high young’s modulus to allow minimumdeflection of the booms, high strength to prevent buckling and low density to give light weight. Inaddition the cost should comply with the unit production cost. It is rarely a case for a material tohave all these qualities. The material selected for the booms is (CES Edupack Software, 2011) epoxycarbonfibre (SMC). This material is strong enough for the booms and less dense than other materialswhich makes it ideal. Moreover the material can handle the extreme expected temperatures of themission, is self extinguishing and is already used for lightweight structures in aerospace. The weightof each boom is approximately 8.4 kg and the cost is AC118. The properties of this material are listedin table 4.18.

Material characteristic Value

Density 1400 kg/m3

Price 14.1 AC/kgYoungs modulus 69000 MPa

Yield strength 221 MPa

Tensile strength 276 MPa

Compressive strength 207 MPa

Minimum service temperature −123 ◦C

Maximum service temperature 166 ◦C

Table 4.18: Epoxy Carbon fibre (SMC)

4.6.3 Boom connection to engine mount

The booms are clamped to the wing by means of steel clamps. The canard booms are attached to thebooms of the engine mount. A steel clamp is used for this connection. One can tighten the clamp tothe desired extent to ensure the booms are well connected.

4.6.4 Finite element analysis on booms

The deflections due to the applied forces on the boom are of vital importance, as it directly affectsthe canard position. An FEM model is made of the boom. Please note that the booms are modeledas a beam with isotropic material properties. However, each layer of carbon fiber should be modeledwith separate elements in the finite element model. This is not pursued due to its complex nature.

The end connecting to the boom of the engine mount is modeled as a clamp and the other end isloaded by the highest load experienced by the canard during a conventional cruise flight regime of theaircraft. The booms are sized, such that these loads do not deflect the boom tip (canard connectionpoint) by more than 40 mm, which results in a canard deflection of 1.0deg. The booms are furthersized to make sure that the worst case scenario loads (not occurring during conventional missions) donot lead to plastic deformation. This is done by ensuring that the maximum stress at the clamped endof the boom is lower than the yield stress of the material. Since this is a relatively simple structurethe mesh is refined to attain a solution with an error of around 3.1%, estimated by CATIA.

Figure 4.20 illustrates the stress distribution of the boom due to loads present during cruise andfigure 4.22 illustrates the stress distribution of the boom due to the worst case scenario of loading.Figure 4.22 confirms that plastic deformation will not occur, as the maximum stress at clamped endof the boom is 159MPa, which is lower than the yield stress, 221MPa of epoxy carbon fiber. Thecondition of the maximum displacement is also met at loads during cruise. Figure 4.21 confirms this.

54 4 Structural Design & Analysis

Figure 4.20: Boom stress distribution due to loads encountered during the cruise flight regime

Figure 4.21: Translational displacements of the boom due to loads faced during cruise

Table 4.19 lists the maximum stresses calculated using the FEM and analytical computations due toloads acting on the boom during cruise and the ’worst case scenario’.

4.7 Control surfaces

The control surfaces needs to be designed and analyzed into more detail. The cross section of thecontrol surfaces was first defined as a triangle with a semicircle connected to it, as is shown in figure4.23a. The triangular section is the actual control surface, the circular section is there to makesure that the flow can transition from fixed parts to movable parts smoothly and also provides someadditional reinforcement to the control surface structure.

Figure 4.22: Boom stress distribution due to worst case scenario loads

4 Structural Design & Analysis 55

Load TypeMaximum Stress [MPa]

Analytical Solution FEM

Cruise 49.5 49.8

Worst Case Scenario 155.6 159

Table 4.19: Comparison of FEM with analytical solution

Figure 4.23: Control surface design. a) Main design; b) Open section shear flows; c) Closed section with

additional circular shear flows; d) Flaperons support.

It should be noted that the design of the flaperons and rudders have the same basic design andwill therefore be treated together. On the other hand, the attachment of the flaperons to the wingbox needs some additional attention, and will be treated seperately in Section 4.7.1.

To perform the structural analysis analytically, a thin-walled cross section is assumed. The pres-sure distribution over the control surface is replaced by a single discrete force. The shear flows arecalculated by first making two cuts, then the first shear flows contributions can be calculated andafterwards the structure is closed again. This is shown in figures 4.23b and 4.23c. Closing the struc-ture means that both of the cells get an additional circular shear flow contribution. To solve thesetwo unknown, two additional equations are required. The first equation states that the twist anglesof both cells should be equal and the second equation states that internal and external forces andstresses should result in moment equilibrium around an arbitrary point. Hence, the shear flows andstresses can be calculated.

The deflection of the control surface due to the pressure distribution was calculated by assumingthat the control surface acts like a thin walled beam and a constant lift distribution is acting over theentire control surface span. Using these simplifications the deflection is calculated.

4.7.1 Attaching the flaperons

The flaperons need to be attached to the rest of the aircraft. This is done by means of hinges onbeams which are connected to the wing box. This is illustrated in figure 4.23d. There is a significantdistance between the hinge point of the flaperons and the sidewall of the wing box. Therefore a stiffand relatively strong connection is required between hinge and wing box. A beam is designed to doso, and using the thin-walled beam theory the beams are sized.

4.7.2 Results

The main sizing parameters as determined by these methods are shown in Table 4.20. The supportbeams have square thin-walled cross-sections: the dimensions in the table indicate the wall widths.

The vertical fins act like a landing gear, which put some extra requirements on the design ofthe vertical fins. Due to time constraints, this has not been designed into further detail, but it isrecommended to do so in the future.

56 4 Structural Design & Analysis

Entity UnitInboard Outboard

RudderFlaperon support

flaperon flaperon Inboard Outboard

Wall thickness [mm] 1.8 1.8 0.75 1.0 0.6

Cross-section dimension [mm] - - - 12 8.0

Table 4.20: Main sizing parameters, determined analytically

UnitMain Engine Propel- Bea- Worm Bevel

Totalshaft shaft ler shaft ring gear gear

Amount [−] 1 1 1 8 1+1 2+2 -

Length [m] 4.62 0.5 0.75 - - - -

Radius [mm] 10 10 10 - - - -

Mass* [kg] 6.5 0.7 1.1 3.0 2.0 1.0 37.37

Cost* [AC] 65.79 7.04 10.66 9.00 15.00 4.00 189.30

Table 4.21: Results of shaft and gear box structural analysis.

4.8 Shafts and gear boxes

The power of the engine in the fuselage needs to be transferred to the proprotors located on the wings.This is done using shafts and gear boxes. To save weight, it is chosen that the gear boxes have a fixedgear ratio. The chosen gear ratios and rotations are shown in Figure 4.24a.

As is shown in Figure 4.24b, in the fuselage a worm gear is used to drive the main shaft. Onboth ends of the main shaft, bevel gears are used to drive to proprotors. The shafts are supported byself-aligning ball bearings, to ensure the friction is kept to a minimum.

The shafts are made of titanium, which is a strong and light material. The main shaft is locatedaft of the wingbox. The propeller shafts intersect the front and back wall of the wingbox. All relevantresults of the structural analysis are shown in Table 4.21. Cost and mass for the gears are indicatedper couple of gears.

Figure 4.24: Shaft and gearbox design. a) Rotations and RPMs; b) Gears and bearings.

Chapter 5

VATOL Simulation

Since the VATOL maneuver is not a common one, a flight dynamics model is made to verify that theFlamingo can actually perform this maneuver. The simulation is performed using a non linear modelprogrammed in Simulink. The simulation uses a linear quadratic regulator (LQR) for control. Howthis simulation works will be discussed in this chapter. The translation between the different framesof reference is discussed in section 5.1. The equations of motion and linearized equations used willbe discussed in section 5.2. The aircraft controller is discussed in section 5.3, the guidance system isdiscussed in 5.4. The simulation is discussed in section 5.5. Finally, the results will be discussed insection 5.6.

5.1 Flight dynamics and quaternions

During the VATOL maneuver the aerodynamic angles α, β and σ can take almost any value between−π

2 rad and π2 rad. This means that the Euler angles conventionally used in flight dynamics cannot

be used as they have a singularity in either α = ±π2 rad, β = ±π

2 rad or σ = ±π2 rad, depending on the

order of the transformation; this is also known as Gimbal lock. The Euler angles can represent anyorientation of any frame of reference but when trying to translate the forces and moments associatedto that frame into another frame of reference, singularities are met when a plane is rotated such thatit coincides with another one (which happens at the angles mentioned above).

~q =

e0

exeyez

(5.1)

A solution to this problem is found in the quaternion representation. A quaternion is a vector withfour inputs which together represent an axis of rotation and the magnitude of the rotation about thataxis (Phillips, 2004). The basic quaternion relation used for the simulation can be found in equation5.1.

In equation 5.1 e0 is a scalar input describing the magnitude of the rotation, and ex, ey, ez describethe so called Euler axis which is the axis about which the rotation takes place1.

1For further information about quaternions and flight dynamics, please refer to (Mooij, 1994) and (Phillips, 2004)

57

58 5 VATOL Simulation

5.2 Derivation of the linear system of equations

In this section the derivation of the linear system of equations from the equations of motion will bediscussed. Since this model is based on quaternions rather than Euler angles, the equations of motionare different from the conventional equations of motion which are based on the Euler angles. Theequations of motion in terms of quaternions can be found in equations 5.2, 5.3 and 5.4 (Phillips, 2004).

Equation 5.2 relates the time rates of change of the velocity components to the change in the di-rection of the gravitational pull, the aerodynamic forces in their respective direction and the roll,pitch and yaw rates.

Equation 5.3 relates the time rates of change of the pitch, roll and yaw rates to the mass moment ofinertia, the angular acceleration, the aerodynamic moments and the angular rates.

uvw

= g

2(exez − eye0)2(eyez + exe0)e2z + e2

0 − e2x − e2

0

+g

W

X + TxbYZ

+

rv − qwpw − ruqu− pv

(5.2)

pqr

=

Ixx 0 −Ixz0 Iyy 0−Ixz 0 Izz

−1 0 −hzb hybhzb 0 −hxb−hyb hxb 0

pqr

+

l + (Iyy − Izz)qr + Ixzpqm+ Txbzbp + (Izz − Ixx)pr + Ixz(r

2 − p2)n+ (Ixx − Iyy)pq − Ixzqr

(5.3)

e0

exeyez

=1

2

−ex −ey −eze0 −ez eyez e0 −ex−ey ex e0

pqr

(5.4)

Equation 5.4 relates the time rate of change of the quaternion to the pitch roll and yaw rates.

For the simulation, these equations need to be represented in a state space formulation of the form

~x = A~x+B~u (5.5)

with vectors ~x =[u v w p q r e0 ex ey ez

]Tand ~u =

[δc δa δe δr δT δ∆T

]T.

Where δT is a change in total thrust and δ∆T is differential thrust setting between the two propellers.

Elaborating the equations of motion gives expressions for u, v, w, p, q, r, e0, ex, ey and ez. Usinga first order Taylor expansion, these equations can be linearized about an equilibrium point. Theselinearizations contain terms that relate to the current state of the UAV and terms that relate to thecontrol inputs (e.g. deflections and thrust settings). The former terms can be found in the A matrix(equation 5.6), the latter ones in the B matrix (equation 5.7).

5 VATOL Simulation 59

A =

CXu 0 0 0 0 0 −CXe0 CXex −CXey CXez0 CYv 0 0 0 0 CYe0 CYex CYey CYez0 0 CZw 0 0 0 CZe0 −CZe0 −CZe0 CZe00 0 0 Cpp Cpq Cpr 0 0 0 00 0 0 Cqp Cqq Cqr 0 0 0 00 0 0 Crp Crq Crr 0 0 0 00 0 0 −Ce0p −Ce0q −Ce0r 0 0 0 00 0 0 −Cexp −Cexq −Cexr 0 0 0 00 0 0 −Ceyp −Ceyq −Ceyr 0 0 0 00 0 0 −Cezp −Cezq −Cezr 0 0 0 0

(5.6)

The linearization of the system of equations results in a decoupling of the system of equations. Thiscan be clearly seen in the A matrix as any rotational state does not affect any translational state andvice versa. This assumption is a valid one as the so-called ”cross coupled” terms are much smallerthan the decoupled ones.

B =

0 0 0 0 CXδT 00 0 0 CYδr 0 0

CZδc 0 CZδe 0 0 00 Clδa 0 0 0 0

Cmδc 0 Cmδe 0 0 00 0 0 Cnδr 0 Cnδ∆T0 0 0 0 0 00 0 0 0 0 00 0 0 0 0 00 0 0 0 0 0

(5.7)

The B matrix is also simplified. Any control surface deflection causes an increased drag, however,this is assumed to be small and therefore set to zero as can be clearly seen in the first row in equation5.7. This prevents the simulation of getting confused, as it could otherwise decide that a change invelocity can also be realised by generating extra drag through a control surface deflection. In theorythis is ofcourse possible, but it makes no physical sense to do so. Any change in velocity should berealized through a change in thrust only. In other words, the B matrix only relates deflections ofcontrol surfaces and engine settings to the rates in which they have a primary effect.

5.3 Aircraft controller

Most aircraft employing an advanced fly-by-wire system make use of a digital controller. A typicalcontroller is a feedback system that takes deviations from a nominal state and transforms them intoa control response to correct for these. Aircraft with relaxed stability can often be made stable byuse of a controller that corrects for the deviations that would otherwise make the state diverge.

The Simulink implementation of the aircraft model uses a linear quadratic controller to control theattitude and translational motion. A feedback controller based on an LQR takes a set of lineardifferential equations, the equations of motion, and minimizes the cost associated with controlling thesystem. The cost is expressed in terms of the deviations of the control mechanisms from their typicalvalues. The aircraft equations of motion derived in section 5.2 can be expressed in the state spaceform, as shown by equation 5.5. In this equation the deviation from the nominal state of the aircraftis expressed as ~x, while its derivative is ~x. The control vector, u, contains the control parameters.

60 5 VATOL Simulation

Figure 5.1: The gains are computed at several instances when using gain scheduling

The appropriate control response to a deviation from the nominal state is determined by equation5.8.

~u = −K~x (5.8)

In this equation K is the control gain. This gain is found by minimizing the control function. Forthe infinite-horizon case, where the time for the system to react is sufficiently large, the cost functionis given as:

J =

∫ ∞0

(~xTQ~x+ ~uTR~u)dt (5.9)

The R- and Q-matrix are the control engineer’s only tools for tuning an LQR-controller. They aresquare matrices with their input variables as inverse squares on the diagonal. According to Bryson’srule (Mooij, 1997), typical values for these parameters are related to the maximum deflections of thecontrol surfaces and typical allowable deviations from the nominal state, respectively. An R-matrixwith high δri values will allow for large deflections of control surfaces. For small δqi values the systemwill react very sudden, and might overshoot the target state when correcting.

The LQR method requires a set of linear differential equations. As the original equations of motion arenon-linear, they need to be linearized, as shown in section 5.2. When linearizing the equations about anominal point, the aircraft is placed in an equilibrium state. This means that as the aircraft deviatesfrom this point, as it does when pitching up, the linearization will become invalid. For this reasonthe aircraft is linearized again at a new nominal point, and a new controller gain is determined. Thecontroller gains can then be interpolated to produce a smooth transition. This method is called gainscheduling and is illustrated in figure 5.1. To realize when the aircraft must be relinearized variousscheduling variables are employed. These are state parameters that are monitored to determine whenthe aircraft has deviated sufficiently from the nominal point to be outside the validity region. Thesevariables should not have the same value twice in the same region considered by the algorithm asthen the guidance will get confused and will think it is in a different part of the maneuver.

5 VATOL Simulation 61

The numerical simulation of the VATOL maneuver is controlled by a controller written in MAT-LAB. It is possible to use native MATLAB functions to determine the gain matrices automaticallyfrom the A, B, Q, and R matrices.

The controller is adjusted to several different flight phases to account for the differences in behaviorexhibited by the aircraft.The aircraft state vector ~x and control vector ~u are are shown in equations 5.10 and 5.11. The controlvector entries have been divided into the relevant deflections. The flaperons serve as both elevatorsas well as ailerons, this functionality has been separated in the vector and is added together to givethe total deflection. The thrust given by the various propellers has also been divided into a generaltranslational thrust and a differential thrust that affects the yaw.

~x =

uvwpqre0

exeyez

(5.10)

~u =

δcδaδeδrδTδ∆T

(5.11)

The gains are generated in a separate function. This is done before the simulation starts. Calculatinga gain involves linearization of the equations of motion around a new point. These linearized equationsthen describe the deviations from this point in which the aircraft is in equilibrium. The gain matrixis produced by using the MATLAB function lqr(A,B,Q,R). Once the gains are determined the controlresponse can be determined through equation 5.8. Here ~x is the error of the state vector with respectto the nominal case. The control vector and the state error is used in equation 5.5 to determinethe change in state vector. To determine the full state this change is multiplied by the time stepand added to the overall state. The state error for the translational motion and rotational rates wasdetermined as follows:

~xerr = ~x− ~xc (5.12)

Where ~xc is the commanded state. The quaternion error, however, cannot be simply subtracted.To determine the error rotation in quaternion form one must divide the current attitude by thecommanded attitude.

~Qerr =~Q

~Qc(5.13)

5.4 Aircraft guidance

The current guidance system is basic guidance system but performs well with the pitch up maneuver.During the pitch up maneuver there are 3 phases in the Simulink model. First there is a trim phase

which lasts 20 seconds. In this phase the aircraft is directed to fly straight ahead with a constant

62 5 VATOL Simulation

velocity. After which a phase starts in which the pitch up maneuver is performed. In this phase theaircraft is commanded to increase its pitch to about 90 degrees while reducing its airspeed. Lastlythe hover phase starts where the aircraft also slowly descents to the target location. Each phase hasits gain which are all calculated before the simulation runs.

During the take off the Guidance system first commands to increase speed while maintaining anear vertical pitch. After reaching screen height the aircraft rotates to reduce the pitch and flies away.

5.5 Simulation

In this section the simulation will be discussed. First the assumptions will be explained in section5.5.1. Next the Simulink model will be discussed in section 5.5.2 and finally the result of Simulinkmodel will be discussed in section 5.6.

5.5.1 Assumptions and aerodynamic data

The main assumptions made by the simulations implementation were:

• Flat, non-rotating earth

• No gyroscopic effects

• Steady atmosphere

• Sea level, standard atmosphere conditions

• Induced velocity is equal to mean induced velocity and is evenly distributed over main wing

The assumption that the induced velocity obtained from the propellers is equal to the mean inducedvelocity is likely to be the assumption with the greatest impact on the accuracy of the simulation. Itwas not possible to accurately determine the velocity profile over the wing in the time span available.However since each propeller covers roughly an entire half-wing, the assumption was accepted.

The aerodynamic characteristics for both models were determined from Tornado citeptornado,and assembled into tables which the simulations could access. All the aerodynamic coefficients werecalculated for a range of angles of attack, pitch rates and canard deflections. This information wasstored in 3D matrices from which the simulations can interpolate the coefficients needed.

The information about the induced velocity from the proprotors was generated using the BEManalysis discussed in section 3.5. These tables were made with as input the propeller thrust andvelocity and as output the induced velocity. Another table was made to calculate the maximumthrust available at different flight speeds.

5.5.2 Simulink implementation

In this section the Simulink implementation of the aircraft model will be discussed. First a generaloverview will be given on how the model works. After that, the three main functions of the modelare discussed namely: the flight model, the guidance system and the controller.

Overview of the Simulink model

In figure 5.2, a top level overview of the Simulink model is shown. The model in Simulink consists ofseveral blocks which are connected to each other. The flight model block outputs the current state tothe bus. This information is fed into the control and guidance systems where the new control surfacedeflection and thrust settings are calculated. Also the visualization is connected to this bus but thisdoes not influence the simulation.

5 VATOL Simulation 63

Figure 5.2: Flight Model

Thrust2

Control Surfac1

Turn Q off

K

To Workspace1

Thrust

To Workspace

controlSubsystem1

Bus

State Vector

State Vector1

V

Subsystem

State Vector

Gain

Velocity

Thrust Right

Thrust Left

Canard

Alleron

Elevator

Rudder

QuaternionDivision

q

rq/r

gain3

Guidance2

Bus1

Thrust Right

Thrust Left

Canard

Alleron

Elevator

Rudder

Figure 5.3: Simulink controller block

Simulink model

In figure 5.4 the Simulink flight model is shown. This model propagates in time and uses the thrustand control settings from the controller as an input. The “6DoF (Quaternion)” block from theSimulink aerospace toolbox calculates the velocities, rotations and position. The output of this blockis used as an input for the next time step.

In the “Effect of Thrust” block the induced velocity is calculated. This is done by entering thecurrent thrust and velocity into a table after which the induced velocity is return. In the “FlightConditions”’ block the current flight conditions are calculated. The dynamic pressure is calculatedby adding up the velocity in the body axis and the induced velocity which only acts in the body xdirection. This induced velocity is used together with the Vbody to calculate the dynamic pressure.

In the “Aerodynamics Coefficients” block the aerodynamic coefficients are looked up from thetables. After which they are fed into the “Forces and Moments” block. Here they are converted tothe forces and moments acting on the UAV.

In figure 5.5 the guidance system for the landing is shown. Using switches the controllers arechosen.

In figure 5.3 the controller is shown. From the bus the current flight conditions are extracted afterwhich the desired state is subtracted. This leaves the error which is multiplied with the gain afterwhich gives the deflections.

Controller

The controller uses the gain matrix supplied by the guidance system and the error in the state. Usingthis information the control surface deflections and thrust settings are calculated as explained in

64 5 VATOL Simulation

Figure 5.4: Flight Model

Gain2

Desired State1

gain 2

K2

gain 1

K

Trim

In1Out1

Out2

Switch6

>= 0

Switch1

>= 0Switch

>= 0

Pitich Up

In1Out1

Out2

Hover and descent

In1Out1

Out2

Bus1

Figure 5.5: simulink guidance block

5 VATOL Simulation 65

!"#!"

$!!"$#!"%!!"%#!"&!!"&#!"

'$(!!" '$$!!" '(!!" '$!!"

!"#$%&

'()*

+(

,-.$/01'($2($/34'$()*+(

Figure 5.6: Altitude versus distance to target during vertical landing

!"!#$"!#%"!#&"!#'"("

(#$"(#%"(#&"(#'"

!" )!" (!!" ()!" $!!" $)!"

!"#$%&'(

)*+&,-'.

/&

0"1+&,2/&

Figure 5.7: Pitch angle versus time during the vertical landing

section 5.3. The maximum deflection of the control surfaces and the rate at which they deflect arelimited. The thrust settings and the rate at which the thrust can change are limited too.

5.6 Results of the simulation

In this section the results of the simulations will be shown. In the first section the landing is discussed,and in the second section the take off will be discussed. From these results it can be concluded thatthe proposed maneuver is feasible.

5.6.1 Pitch-up maneuver

The results of the simulation in the Simulink model look very promising about the stability duringthe pitch up maneuver and landing. In figure 5.6 the distance to target and altitude position areshown over time. This graph clearly shows the maneuver the Flamingo makes when it lands at itsdestination. In figure 5.7 the pitch angle is shown versus time. As can be seen the pitch remainsconstant during hover however is not exactly vertical, but approximately 89o. It should be notedthat the aircraft can be guided to the target location by slightly changing the pitch angle. Lastly infigure 5.8 the absolute velocity is shown over time. This figure shows clearly shows that the pitch upmaneuver starts with changing controller as the aircraft first pitches down. As soon as the aircraftpitches up, it loses speed due to the gravitational pull even though the propellors provide more thrust.It should be noted that the graph shows the absolute velocity in the earth reference frame this causesthe small bumps when the velocity nears zero.

66 5 VATOL Simulation

!"

#!"

$!"

%!"

&!"

'!"

(!"

!" '!" #!!" #'!" $!!" $'!" %!!"

!"#$%&'()&*)+

",-'./0&-"%1$*

)23456)

7&3")256)

Figure 5.8: Velocity versus time during the vertical landing

0

50

100

150

200

250

-­‐500 0 500 1000 1500 2000

Altitude  [m]  

Distance  from  target  [m]  

Altitude  versus  distance  

Figure 5.9: Altitude versus horizontal distance from target during the vertical take off

5.6.2 Vertical take off

Also the simulation off the vertical take off works well. Shown in figure 5.9 is the X location versusthe Z location. In figures 5.10 and 5.11 the pitch and absolute velocity are plotted versus time. Theinitial bumps can be explained by the fact that, as the aircraft is tipping over, it does not have enoughspeed to generate enough lift to counter the pitch down moment and the controller has to aggressivelycorrect here. The second bump in the plot betrays the switch of gains in the climb phase.

5 VATOL Simulation 67

!"

!#$"

!#%"

!#&"

!#'"

("

(#$"

(#%"

(#&"

(#'"

!" (!" $!" )!" %!" *!" &!"

!"#$%&'(

)*+&,-'.

/&

0"1+&,2/&

!"#$%&3+-242&51+&

Figure 5.10: Pitch angle versus time during the the vertical take off

0

10

20

30

40

50

60

70

0 10 20 30 40 50 60

Velocity  in  Xearth-­‐direcon

 [m/s]

Time  [s]

Figure 5.11: Velocity versus time during the vertical take off

68 5 VATOL Simulation

Chapter 6

Subsystem Design

In this chapter the subsystems will be discussed. The systems discussed in this chapter are; the powersystem in section 6.1 this is followed by the electrical block diagram giving an overview of all theelectrical systems in section 6.2, the control and navigation system is discussed in section 6.3, thecommunication system in section 6.4, the data handling in section 6.5 and finally the support systemdiscussing everything that is not on the Flamingo itself, will be discussed in section 6.6.

6.1 Power system design

In the powers system design, three different components have been considered: the engine, the elec-trical power system and the fuel tank and tubing.

The engine chosen is the Mistral G-200. This is a 1960cc, 200hp wankel engine that runs on avgas andmogas(The rotary engine for 21st century general aviation G-200, n.d.). This engine has been chosenfor its relatively small dimensions, its reliability and its performance in terms of fuel consumption athigh altitudes.

The electrical power system consists of a battery, actuators for the control surfaces, wiring and thesensor package. The electricity is generated by the generator which comes with the engine. It is thenfed to the battery which feeds it to the aircraft. The battery is sized such that it can supply all theelectrical systems of the aircraft during the hover maneuver and for a duration of 15 minutes at fullcapacity after complete failure of the engine.

The three aspects that determined the design of the fuel system the most are the volume of the tankthe position in the aircraft and the fact that for the VATOL sloshing should be kept at a minimum.

A tank volume of 60 liters is chosen. This allows for extended missions up to 300km requiring53l and leaves sufficient volume for loitering. For redundancy reasons it is chosen to divide this overtwo tanks which both have separate tubing connecting the tanks to the engine. The design allowsthe aircraft to fly back with cargo if a fuel pump (or part for the fuel system) fails during any part ofthe mission except for during the VATOL maneuver.

The tanks are placed in the fuselage next to the engine. Placing the tanks in the wings has beenconsidered, however, there is not enough space there due to the large control surfaces and the shaftsthat drive the propellers. Placing the fuel tanks close to the center of gravity improves maneuverabilityas the mass moment of inertia about the Xb and Zb axis reduces.

Sloshing is an important consideration for this aircraft. During the VATOL manouver large dis-placements of fuel could lead to stability issues and fuel might no longer be able to reach the engine.To solve this problem, a special type of fuel tank is designed. This tank with triangular cross section

69

70 6 Subsystem Design

Generator

Charge circuit

Battery

Onboard Flight Computer

Payload Release System

Control & Navigation System

Canard actuatorsCommunication

system

Backup battery

Regulator

Figure 6.1: Electrical Block Diagram

consists of a bladder divided into compartments1. This bladder is glued to two stiff plates on theadjacent side and the hypotaneous of the triangle. As there will be no air in the bladder the tankshrinks as fuel is consumed. The fuel is distributed evenly over the tank by the stiff plates, even inhovering flight and in soing so, reducing sloshing.

Table 6.1 shows the different components and their weights.

component mass [kg]

engine 132

battery and wiring 10

actuators 10

fuel tank and tubes 10

Table 6.1: Different components and their masses

6.2 Electrical block diagram

The electrical block diagram gives a clear overview of the electrical components inside the UAV. Thecomponents that are generating or using electrical power are listed in blocks, the arrows representthe connections between the different subsystems.

For the UAV the electrical block diagram is shown in Figure 6.1. The heart of the diagramis the electronic regulator. This unit distributes the available power to the right places. Usuallyelectrical power will come from the generator. The generator is integrated in the engine and convertsmechanical power into electrical power. The electricity out of the generator is than used to chargea battery through a charger circuit. In case the battery is fully charged and the generator producesenough electrical power, the energy can be distributed to the subsystems directly. If however thegenerator is not providing enough power, e.g. when all power of the engine is required for hovering,the subsystems use power of the battery. As soon as less power is required for propulsion, the batteryis charged again.

Furthermore a backup battery is present in case the main power supply fails. This battery willnot be attached to the charge system to reduce the probability that both batteries are damaged dueto a failure in the charger. The backup battery has a limited capacity and can only supply electric

1Similar to an air mattress.

6 Subsystem Design 71

power for a short time period, thus the UAV will have to perform an emergency landing as soon aspossible after failure.

6.3 Control & navigation system

The control & navigation system is very important for the UAV, as the vehicle has to perform itsmission autonomous. The control & navigation systems of the UAV consist of a chip set and additionalsensors to find an appropriate landing site and to monitor the ground during landing. They will betreated separately in the following sections.

6.3.1 Autopilot chip set

The MicroPilot MP21283X has been selected as the UAV autopilot. Previous autopilots by the samemanufacturer have already been used by NASA and MASS. The latter one has even used it for a tailsitter UAV. Therefore this chip set seems a good choice (MicroPilot MP2128, 14-06-2011). The mainadvantages of the chip set are listed below.

1. triple redundancy: three independently operating auto-pilots

2. 3 gyros: pitch, roll and yaw rate measurements for attitude determination

3. integrated GPS: used for global positioning

4. multiple communication links: e.g. radio communication

5. lightweight and small: the mass is only 28 g and it is 10 x 4 cm in dimension

6.3.2 Short distance landing sensor

For the depth sensing during landing, PrimeSensor technology is used(PrimeSense: Motion ControlBeyond the Kinect, 16-06-2011). It projects a grid of infrared light on the surroundings, and thereflections are observed by a infrared sensing camera. The system is also equipped with a conventionalcamera. The sensor is able to see depth during day and night within a reasonable distance.

6.3.3 Long distance landing sensor

For landing site scanning of the ground the RIEGL LMS-Q160 laser radar is used. It is lightweightand especially designed for UAVs (RIEGL LMS-Q160, 14-06-2011). It enables the Flamingo to makea detailed scan of the surrounding environment and identify human beings on the landing site.

6.4 Communications

The Flamingo has a need to communicate with the operators in order to make sure missions goaccording to plan or to change the mission during flight. The range requirement of 200 km limits theoptions available as most communication systems have a far shorter range.

A communication system based on broadband satellite Internet is chosen. This is done for thefollowing reasons. First the Internet connection gives great flexibility on the location of the operator.Second satellite Internet connection works almost everywhere as shown in figure 6.2. For the antennathe Explorer 727 is chosen and is shown in figure 6.3 . This is a small dome about 46 cm in diameterin which the antenna can continuously track the satellite. The antenna is mounted on the rear of theaircraft behind the engine.

This system is augmented with a short range relay system in which the UAVs can send data viaeach other to base of operation. This system uses a Ku band antenna with a line of sight range ofabout 50km. This means that for every mission distance increase of 50km, starting from 50km, at

72 6 Subsystem Design

least one extra Flamingo is needed. This relay system is preferred for data transfer as the satelliteInternet connection is more expensive for sending data.

When the aircraft is hovering the satellite antenna is not pointed towards the air and thereforecan not connect to the satellite. If communications are deemed necessary, the Flamingo can send itssignal via the Ku band antenna to a nearby Flamingo which in turn can either relay the data eitherto the satellite or through the relay system.

Figure 6.2: Satellite Internet coverage of the BGAN net-work Figure 6.3: Satellite Antenna

6.5 Data handling

In this section the data handling is explained in detail. The data handling subsystem carries andstores data between various units and the ground segment. For the concerned mission an on boardcamera takes pictures of the surroundings, which are needed to determine a suitable landing spot. Asuitable camera for this mission is a CCD camera (charged couple device), because it produces highquality image data. The data handling block diagram is illustrated in figure 6.4. The camera takespictures and the transceiver of the system will transmit the images to the ground station, so thatthe images can be displayed on a monitor. The camera with two degrees of freedom is set up on aservo platform which provides control of a desired operation through the use of feedback. Throughnavigation the UAV knows its attitude and location. This information is transferred to on boardcomputer memory. The driver of the servo platform will download this file and determine the relationbetween the body frame coordinates of UAV and line of sight of the camera. Once a suitable landinglocation is detected, the servo moves the camera to pin point to the target location by using the twodegrees of freedom movement.

6.6 Support system

One single UAV is always a component in a larger system necessary to successfully perform a mission.This includes control stations, equipment and tooling, communication links, and other hardware.Crew requirements are considered separately as a part of the operations and logistics, in chapter 7.The support system is designed to manage 10 aircraft at any given time. To do this the control stationneeds to be able to control 10 aircraft simultaneously. This puts high demands on the design so asto ensure sufficient overview, while also providing details on each aircraft, for the pilot. A detaileddesign of the support system is beyond the scope of this report, but a general description will begiven.

Being the most critical component of the support system the actual control system is doubleredundant. This is achieved by having a pilot and copilot operate independent systems that each can

6 Subsystem Design 73

Figure 6.4: Data Handling Block Diagram

operate the entire fleet independently. Under normal conditions direct pilot control is not necessary,only flight plan uploading and monitoring. Each of the control systems consists of three screensproviding sensor data and imagery from selected UAVs, as well as flight data. Pilot control is achievedby inserting the destination after which the UAV will fly to the destination autonomously.

Spare parts and tooling has also been considered a part of the support system. Parts that havea higher risk of damage, or are more susceptible to wear, have replacements brought to site with thesystem, so as to reduce possible downtime in an emergency situation. Carbon-fibre and resin is alsobrought for field repair of non-structural composite parts.

The following list summarizes the contents of the support system:

Control system components

• Air conditioned tent

• Diesel-powered generator

• Antenna rack

• Computer rack with chair

• Human interface

Tooling

• Rivet gun for cargo pods

• General tools for mainentance

• Field lay-up tools

• Jig for attaching cargo

Spare parts

• Propellers

• Landing gear

• Booms

• Assorted carbon fiber and resin

74 6 Subsystem Design

Chapter 7

Operations, Logistics & SystemEngineering

In this chapter, some relevant topics not directly concerning the technical design UAV itself aretreated. In section 7.1, the operations & logistics of the UAS will be explained. Then, in section 7.2the production plan is discussed. A cost estimation is performed in section 7.3. In section 7.4 theoperational return is elaborated. Thereafter the risk map of the UAS is shown in section 7.5. Then,the RAMS characteristics of the UAS are described in section 7.6. Finally, the FFD and FBS areshown in section 7.7. Future project design and logic is shown in 7.8.

7.1 Operations & logistics

In order to meet the requirements of the customer, the flow of goods and services from the pointof operation to the point of delivery needs to be managed. The logistics involve a combination oftransportation, material handling, warehousing, stocking, packaging and sometimes also security. Thedetails of operations and logistics are provided in the section below.

7.1.1 Transportation

The aspect of transportation involves not only the transport of UAVs, personnel and support equip-ment to the site before the mission and taking it back when the mission completes but it also involvesthe transportation of payload to the UAVs from the nearest hangars or warehouses if available.

Transportation in A400M

For emergency relief missions, the UAS is first brought to the airport that is closest to the disaster site.An A400M is used for transportation. One of the requirements is that ten UAVs, support equipmentand support staff should be able to fit in this A400M. Therefore the UAV has been designed suchthat some components can be disassembled rather easily. The way individual components are putin A400M is illustrated in figure 7.1. The cargo bay has a width of 4m, its length is 17.7m and theheight is 4.2m. The presence of the cargo door gives additional space which is also used.

The booms, vertical fin, rotor hub and the rotor blades are detachable. After these componentsare removed from the aircraft, five aircraft are placed on top each other in a rack. The height ofone rack containing 5 aircraft amounts to 3.6m, leaving a 20cm clearance between the top of theaircraft and the roof of the A400M. Ten canards are placed on top of each other. The booms arestacked in another rack by pair, one rack supporting ten booms. There are 20 vertical fins which arestacked as shown in figure 7.1. During each sortie the cargo pod is left behind at the disaster area. Toprovide aid for three days, there are 180 cargo pods needed (ten UAVs, each performing 6 missions

75

76 7 Operations, Logistics & System Engineering

Figure 7.1: A400M layout where ten UAVs are loaded with support staff and support equipment

per day during three days). To fit this amount of cargo pods in A400M, the cargo pods are foldedin different ways and stored at different locations in the cargo bay. The placement of the cargo podscan again be found in 7.1. Number of personnel needed is 15 and these 15 people and their luggagesare also transported by A400M. One person’s luggage’s length, height and depth are 70cm x 40cm x40cm, respectively. In A400M there is also enough space given for the support equipments and thesesupport equipments are explained in section 6.6. During the transportation it is made sure that allthe components are fitted in such a way that they do not move or fall. When the mission is completeall the UAVs and support equipment are disassembled and transported back in the same manner. Itis possible to take avgas for the missions. There is sufficient space allocated to store fuel for the 180missions.

Transportation of payload

It is assumed that a warehouse or hangar will be present in the vicinity of the airport and theorganization for which the mission is performed will have ensured that there is always a continuoussupply of commodities. The payload is then transported by either cranes or forklifts to the UASstation from where it can be loaded in the pods and attached to the UAV.

7.1.2 Personnel

Trained personnel is needed to operate UAVs in order to accomplish the mission. The amount ofpersonnell as wel as their required capabilities and expertise are elaborated in this section.

Amount of personnel

One can determine the amount of personnel needed by calculating the operational hours of a UAVand the work hours per person. The number of people needed are calculated as follows. Total missiontime of one UAV is calculated in table 3.6 in section 3.7 and that is equal to 2.49hr. The UAV isperforming six sorties per day. To control the ten UAVs during the mission time of six sorties requiresthree people. Except to control the UAVs there are people needed to attach the payload and for themaintenance. By taking a duration of fifteen minutes to attach the payload and one hour for themaintenance per sortie this amounts to 75 man hours (1.25hrs times 6sorties times 10UAV s). Ifone person works eight hours per day, ten people can perform the tasks of attaching payload andmaintenance of the aircraft (75

8 = 9.375). There is also personnel needed to perform other tasks, likeadministration work. It is assumed that two people are enough for these tasks. There are in totalfifteen people needed on board.

7 Operations, Logistics & System Engineering 77

Training and experience

The UAS personnel needs to have an appropriate training to accurately and safely operate a UAS. TheUAS personnel should be able to perform several different kinds of jobs which include coordinatingand planning ground and flight operations and executing them, briefing and debriefing, managingresources and ensuring ground and flight safety.

7.1.3 Material handling

Material handling involves storage, control and protection of materials through all the processes ofmanufacturing, consumption and disposal.

Storage

Storage of materials, before a UAS is send to a disaster area, is usually done in warehouses. Materialis received in the warehouse, inspected, stored and finally dispatched to the site when needed. It isassumed that there will be no warehouse available in the vicinity of the airport. Storage at the airportwill either take place in an available hangar at the airport or in tents which are to be brought withthe A400M. Moreover, spare parts and necessary maintenance equipment like lubricants also need tobe stored which might be done in the same hangars or tents.

Disposal of Payload containers

During operation, the cargo pods are left behind at the disaster site. Since mission needs to befulfilled in an environment friendly manner, the pods will be made of white spruce, which is a naturalmaterial and can safely be burned by the people on site. Choosing a wooden cargo pod ensures thatthere is no toxic waste as a result of delivering goods and that there will be no toxic fumes createdwhen burning the pod.

7.2 Production plan

This section will discuss which manufacturing methods will be used to produce the Flamingo as wellas the on site assembly.

7.2.1 Manufacturing

The material used for most of the structural parts is aluminum 7075. First, the wing box will bediscussed. The wing box consists of four plates, four spar caps, stringers and ribs. The plates havean increasing width when looking from the wing tip to the root. This can be manufactured by rollingwhere a block of aluminum is passed through a pair of rolls, which are moved towards each otherto obtain a decreased thickness. Because the plate has a constant width the plate has to be cutafterwards to obtain a tapered plate. Since the spar caps also have an increasing width, it can bemanufactured in the same way. However, since the spar caps have an L-shape they also need to bebent afterwards. The stringers are not tapered and have a common geometry so they will come offthe shelf. The ribs can also be made using rolling only the edges need to be cut at the end in such away that the ribs fit into the wing box. There are several booms used in this structure. Booms areused to connect the engine mounting and the landing gear with the front and rear spar. Since theseparate booms have a simple shape they will be off the shelf as well.The booms that connect the canard to the main wing are made form carbon fiber composites. Thesebooms are generally available in a wide range of diameters and thicknesses and will come off the shelfas well.

Since the wing box is the primary structural component the wing box will be assembled first.Thereafter the shaft and gear box will be located along the wing box. Then, the skin can be placed

78 7 Operations, Logistics & System Engineering

on the wing after which the fins and the ailerons can be connected. Next, the engine mount andthe support booms of the canard, cargo pod and landing gear can be assembled. As one of the lastcomponents the engine will be assembled. Finally, the last part of the assembly is to connect thecanard and the propellers.

7.2.2 On site assembly

On the operating airport, the Flamingos are to be assembles before operation. First the two boomswill be connected to the fuselage after which the canard can be attached. Now, the UAV is ready foroperation and can thus be fueled and loaded. In order to prepare the payload, the cargo box mustbe unfolded, riveted and loaded with relief. It can be attached to the aircraft using a forklift or jig.

7.3 Cost break-down structure

Requirement D.10 states that the unit production cost of the Flamingo should be less than AC 250000.Requirement D.9 states that the direct operating cost should be less than AC500 per sortie. Fromstructural analysis and chosen subparts it is now possible to estimate the unit production cost. Insection 7.3.1 the unit production cost is given. In section 7.3.2 the direct operating cost is computed.

7.3.1 Unit production cost

The unit production cost is given in table 7.1. The cost that is calculated for the structural parts ofthe aircraft is the cost of the material that is used. For example, for the wing, the cost of the materialthat is needed to build the wing box is calculated. It is important to note that the manufacturing costof these parts is not included as they are difficult to approximate and also depend on the locationof production. The cost of the material per kilogram is found from CES edupack (CES EdupackSoftware, 2011). From the detailed CATIA drawing of the wing box from cross station C (see figure

Aircraft component Cost [AC] Reference

Engine 27000

Wing box 67

wing box of canard 6

Landing gear 756 AirSuppliers (15-06-2011)

Two booms 236

Vertical fins ...

Gears and bearings shafts 190

satellite communications 15858 Ground Control (15-06-2011)

computers for aircraft 200

Lidar 30000 Aplique Kit (16-06-2011)

Primesense sensors [5x] 530 PrimeSense: Motion Control Beyond the Kinect (16-06-2011)

Auto pilot: MP21283x 9780 MicroPilot MP2128 (14-06-2011)

Electrical sub system 500

deployment mechanism system 500

Unit cost 86000

Table 7.1: Flamingo unit cost

4.1) to the wing tip a weight was found of 17.2 kg. Based on this value an estimation of the totalweight of the wing box was made which equals 50kg. Since the price per kilogram for the aluminumused is 1.33 AC/kg the total price for the aluminum required for the wing box equals AC67. Thematerial cost of the wing box of the canard is unrealistic. Since the weight of the wing box of thecanard was 4.1kg (see section 4.2), it gives us low material cost. The wing box of the vertical fins are

7 Operations, Logistics & System Engineering 79

not calculated. Since the cost of the wing box of the wing and the canard look unrealistic, it is hardto give an estimation for the cost of the wing box of the vertical fins. The cost of the deploymentmechanism that is given is an estimation, since the references for the deployment mechanism werehard to find. It should be noted that the low value for the unit production cost does not mean thatthe requirement D.10 is met. Since the cost, like manufacturing cost, overhead cost, waste materialcost are hard to estimate, they are not included.

Since the UAV is controlled from the ground, there should be a support system on the ground.The cost of the support system is estimated to be AC36996 (Hardy Diesel Generators, 16-06-2011).

7.3.2 Direct operating cost

In table 7.2 the direct operating cost is given. The direct operating cost calculation can be found insection 3.7. From table 7.2, it can be seen that the total direct operating cost per sortie is less than

Aircraft component Cost [AC]

fuel 100

cargo pod 36

communication 15

Total 151

Table 7.2: Direct operating cost

AC 500 which is stated by the requirement D.9 and it means that the requirement is met. It must bebeard in mind that these costs do not include maintanance and crew salaries.

7.4 Operational benefit

This section compares the costs involved in performing a conventional mission of delivering a payloadof 250kg to a range of 200km using the designed UAV and the benchmark helicopter Robinson R-44.The operational return attained by the user are then estimated. Please note that an in depth anal-ysis of maintenance costs and long term operating costs are not considered due to lack of time andreference data. The communication costs are assumed equal for the Flamingo and helicopter and aretherefore not taken into account.

Section 3.7 estimated the operational cost of the Flamingo UAV without communications to be AC136.This cost is composed of a fuel cost of AC100 and the cargo pod cost of AC36. Section 2.6 estimatesthe fuel used by the Robinson R-44 to be around 148l. At an Avgas fuel price of 2.74AC/liter 1, thisresults in an operative cost of AC405. It is assumed that this helicopter does not require an externalcargo pod that will be left at site. Please note that the costs of personnel required for the operationof the UAV and the cost of the helicopter pilot have not been taken in to consideration due to lackof reference data.

The UAV results in a lower operational cost of AC269 compared to the benchmark helicopter, a 66%reduction. However, it must be noted that this benefit only applies in comparisons with helicoptershaving a similar payload. Hence, the designed UAV will be most beneficial when a relatively smallpayload is to be delivered to numerous areas, whereas a helicopter excels in delivering a large payloadto one location.

1Rotterdam airport, 27-3-2011

80 7 Operations, Logistics & System Engineering

7.5 Risk map

A risk map has been composed for the detailed UAV design. The risk map is a convenient tool toperform risk analysis and mitigation. The distinguished types of failure are listed below. The riskmap is shown in table 7.3.

1. Single engine failure: the UAV design first consisted of two engines at the wings, but now it onlyconsists of a single engine in the fuselage. The risk was mitigated this way, because having two enginesdid not improve the consequence of a single engine failure, while the probability of failure worsened dueto having two engines.

2. Landing gear failure: the conventional landing gear is retractable for aerodynamic reasons. Duringvertical landing, the UAV lands on the cargo pod. When the conventional landing gear fails to go down,the UAV is still able to land vertically on the cargo pod. If the cargo pod is already released and thelanding gear fails to go down, then an emergency landing has to be performed. Landing conventionallyis not an option as this will damage the propellers and the torque rods in the wingbox and potentionallyeven the engine. In this case the aircraft should therefore make a VATOL maneuver again and land onits fins. The worst case scenario is now that the control surfaces will be damaged.

3. Payload release failure: if the payload release mechanism fails, the UAV might not be able to get backto the airport because in certain cases the UAV will not be able to take off. This risk can be mittigatedby instaling a check mechanism that allows to check if the release mechanism works before the VATOLmaneuver is initiated.

4. Stability due to wind and wind gusts: during VATOL, computer power is necessary to keep theUAV stable. However, the autopilot chip set is triple redundant.

5. Wing or proprotor blade breaking: When a wing or a proprotor blade (partly) breaks, for exampledue to supersonic tip speeds, it is likely that more parts break of due to heavy vibrations. Calculatingthese vibrations is beyond the scope of the project and has not been looked into.

6. Failure due to debris on site: the chance of debris getting sucked into the proprotors is quite high onrough terrains. However the chance that this debris is large and heavy enough to significantly damagethe UAV is extremely small. This only has consequences for the maintenance rate and not so much forcomplete loss of the system.

7. Failure due to bird strike: bird strikes are quite likely to occur from time to time, however theinvolved consequence is marginal. Risk mitigation can be done by chasing away birds using sounds. Thenoise of the proprotors will probably frighten off birds in the neighborhood.

8. Injuring people on site: the proprotors are not shielded, so collision with humans might be lethal.Risk mitigation can be performed by using visual and sound signals, which will notify the people of thedanger. Also the noise of the proprotors itself might scare them off.

9. Damage to the UAV due to people on site: the vertical landing, payload release and vertical take offdo not require people near the aircraft. The extensive sensor system makes sure that the Flamingo onlyperforms the landing maneuver if there is absolutely no human being on the landing site. Additionally,people might already be scared off by the measures as described before. Therefore the likelihood ofpeople damaging the UAV is rather small.

7.6 RAMS analysis

In this section the Reliability, Availability, Maintainability and Safety (RAMS) will be discussed. Theanalysis of the RAMS characteristics are important for mission design and to get an idea as to whatcould be considered ‘proper use’ of the UAV.

7 Operations, Logistics & System Engineering 81

Lik

elih

ood

of

failu

re→

Feasiblein theory

Working laboratorymodel

Based on existingnon-flight engineering

3

Extrapolated fromexisting flight design

Proven flightdesign

2,6,7,8 1,9 4,5

Negligible Marginal Critical Catastrophic

Consequence of failure →

Table 7.3: Risk map

7.6.1 Reliability

This UAV uses a single 200hp Wankel engine as a powerplant. This means that single engine failureis critical. To mitigate this risk, very strict reliability demands have been put on the powerplant. Asthe Wankel engine has very few (and relatively slow) moving parts, the risk of failure is far lower forthis engine type than a piston engine. Additionally if a Wankel engine loses compression, overheats,or experiences any other form of typical engine failure it will still continue to produce a reducedamount of power. Since the fuel system is split in two independant systems, the aircraft will be ableto continue to fly at reduced power setting on the remaining fuel in the working system. This meansthat, even if one of the pumps fails the aircraft can still make it back safely. Imbalance of the aircraftdue to this unsymmetric fuel tank usage can be counteracted up by the flight controller with use ofthe control surfaces.

7.6.2 Availability and maintainability

The engine requires limited checking and maintenance, comparable to a small aircraft. Especiallybecause the engine is not running at maximum power during a large part of the mission. The actualmaintenance man hours per flight hour (MMH/FH) will have to be determined in a later stage of thedesign process.

7.6.3 Safety

UAVs are inherently safer than manned aircraft in the sense that they do not carry people on board.However, since this aircraft will have to land in an uncontrolled environment, there might be nonspecialized people around who can be harmed during payload deployment. Especially since thepropellers are relatively close to the ground, people standing too close to the aircraft might getinjured. It is therefore very important to only select landing sites where bystanders can move awayeasily. An accident like that will most likely also damage the UAV severely as it is most vulnerableduring the VATOL maneuver.

7.7 Functional flow diagram & functional breakdown structure

To illustrate the tasks required by the system to perform a mission, a Functional Flow Diagram (FFD)and Functional Breakdown Structure (FBS) are made. These are shown in figure 7.3 and figure 7.2.

The FBS shows the general tasks which have to be performed during a humanitarian relief mission.It therefore includes the loading of the A400M and the transport to the operation base. Extra emphasis

82 7 Operations, Logistics & System Engineering

is put on the delivery of the payload because that is the phase where the UAS differs from conventionalmethods of providing emergency relief.

The FFD indicates the order in which the tasks of the ‘operate UAS’-block of the FBS must beperformed into much more detail. This also includes the more technical steps of the procedure andtherefore also shows the technical characteristics of the Flamingo.

7.8 Future project design and logic

In figure 7.4 a Gantt chart is shown on what the future project might look like. First the aerody-namic and structural design needs to be finished. With the preliminary results some parts may bemanufactured and a prototype should be built which can be used for testing and certification. Theentire development time is estimated to be little over 2 years. After which the Flamingos can be builtand sold.

7 Operations, Logistics & System Engineering 83

Op

erat

e th

e Fl

amin

go

Tran

spo

rt t

he

UA

S to

th

e o

per

atin

g b

ase

Pre

par

e U

AS

for

op

erat

ion

Op

erat

e U

AS

Load

th

e A

40

0M

w

ith

th

e U

AS

Fly

the

A4

00

M t

o

the

op

erat

ion

bas

e

Ass

emb

le U

AV

s

Set

up

su

pp

ort

sy

stem

Pre

par

e ca

rgo

po

ds

Fuel

an

d lo

ad t

he

UA

Vs

Per

form

flig

ht

man

euve

rs t

o t

he

dis

aste

r si

te

Tran

smit

sta

tus

to

sup

po

rt s

yste

m

Del

iver

pay

load

Per

form

flig

ht

man

euve

rs t

o

op

erat

ing

bas

e

Ref

uel

an

d r

elo

ad

UA

V

Per

form

m

ain

tan

ance

Tran

smit

sta

tus

to

sup

po

rt s

yste

m

Iden

tify

lan

din

g si

te

usi

ng

LID

AR

Per

form

ver

tica

l at

titu

de

lan

din

g m

aneu

ver

Rel

ease

pay

load

Per

form

ver

tica

l at

titu

de

take

off

m

aneu

ver

Figure 7.2: Functional Breakdown Structure of a typical emergency relief mission

84 7 Operations, Logistics & System Engineering

Co

nd

uct p

re-

flig

ht

pro

ce

du

res

Initia

lize

syste

ms

Ta

ke

off

Cru

ise

to

de

stin

atio

n

La

nd

ve

rtic

ally

Un

loa

d

ca

rgo

Ta

ke

off

ve

rtic

ally

Cru

ise

ba

ck

to a

irp

ort

La

nd

at

airp

ort

Pe

rfo

rma

po

st-

flig

ht

pro

ce

du

res

Co

nd

uct p

re-f

ligh

t

pro

ce

du

res

Insp

ect a

ircra

ft

Initia

lize

syste

ms

Ta

ke

off

La

nd

ve

rtic

ally

Un

loa

d c

arg

o

Ta

ke

off v

ert

ica

lly

Ch

eck o

il a

nd

co

ola

nt le

ve

lsR

efu

el a

ircra

ft

Sta

rt e

ng

ine

Initia

lize

co

mp

ute

r

syste

ms

Ch

eck p

ow

er

leve

ls

Te

st

co

mm

un

ica

tio

n

link

Ta

xi to

ru

nw

ay

Re

qu

est

pe

rmis

sio

n to

ta

ke

off

Ap

ply

fu

ll e

ng

ine

po

we

r

Ro

tate

ca

na

rd

an

d fla

pe

ron

s to

pitch

up

Clim

b to

cru

ise

altitu

de

Ap

ply

fu

ll e

ng

ine

po

we

r

Ro

tate

ca

na

rd to

pitch

up

un

til

ve

rtic

al a

ttitu

de

an

d z

ero

ve

locity

Re

du

ce

pro

pe

ller

pitch

to

de

sce

nd

un

til to

uch

do

wn

Re

lea

se

pa

ylo

ad

locks

Re

du

ce

pro

pe

ller

pitch

to

pro

du

ce

0.5

*ma

xT

hru

st

Incre

ase

pro

pe

ller

pitch

un

til lif

t o

ff is

ach

ieve

d

Ro

tate

fla

pe

ron

s

afte

r re

ach

ing

scre

en

he

igh

t to

pitch

do

wn

Asse

mb

le c

arg

o

po

d

Atta

ch

ca

rgo

po

d

to a

ircra

ft

Use

ca

me

ras a

nd

LID

AR

to

id

en

tify

lan

din

g s

po

t

Re

du

ce

pro

pe

ller

pitch

an

d r

ota

te

ca

na

rd to

re

du

ce

altitu

de

Clim

b to

cru

ise

altitu

de

La

nd

at a

irp

ort

Re

ce

ive

pe

rmis

sio

n to

la

nd

Re

du

ce

pro

pe

ller

pitch

, ro

tate

ca

na

rd to

re

du

ce

altitu

de

De

fle

ct fla

pe

ron

s

Re

du

ce

pro

pe

ller

pitch

to

bra

ke

airp

lan

e

Pe

rfo

rm p

ost-

flig

ht

pro

ce

du

res

Insp

ect a

ircra

ft fo

r

da

ma

ge

Ta

xi o

ff r

un

wa

y

1.0

2.0

3.0

4.0

5.0

6.0

7.0

8.0

9.0

10

.0

Re

f. 1

.0

Re

f. 2

.0

Re

f. 3

.0

Re

f. 5

.0

Re

f. 6

.0

Re

f. 7

.0

Re

f. 9

.0

Re

f. 1

0.0

1.1

1.2

1.3

1.4

1.5

1.6

2.1

2.2

2.3

3.1

3.2

3.3

3.4

3.5

5.1

5.2

5.3

5.4

5.5

5.6

6.1 7.1

7.2

7.3

9.1

9.2

9.3

9.4

9.5

10

.1

Figure 7.3: Functional Flow Diagram of a typical emergency relief mission

7 Operations, Logistics & System Engineering 85

ID Task

Mode

Task Name

1 Design and development

2 Perform detailed aerodynamic design

3 Perform detailed structural design

4 Design flight software

5 Manufacturing

6 Manufacture parts

7 Assemble prototype

8 Manufacture UAV

9 Testing

10 Perform static testing

11 Perform flight testing

12 Perform hover testing

13 Certification

14 Perform certification

15 Perform marketing

16 Perform detailed market analysis

17 Perform marketing

18 Sell UAV

19 Operation

20 Operate UAV

21-3 30-5 8-8 17-1026-12 5-3 14-5 23-7 1-1010-1218-2 29-4 8-7 16-925-11 3-2 14-4 23-6 1-9 10-1119-1 30-3 8-6 17-826-10 4-1 14-3

21 March 11 August 1 January 21 May 11 October1 March 21 July 11 December1 May 21 September11 February1 July 21 November

Task

Split

Milestone

Summary

Project Summary

External Tasks

External Milestone

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only

Manual Summary Rollup

Manual Summary

Start-only

Finish-only

Deadline

Progress

Page 1

Project: schedule

Date: Fri 17-6-11

Figure 7.4: Gantt chart for further development

86 7 Operations, Logistics & System Engineering

Chapter 8

Conclusions

This chapter concludes the DSE report of group 6 on the design of a UAV for emergency relief pur-poses. During this project, the Flamingo came to life. After two weeks of system engineering andrequirement analysis; initial designing of the preliminary concepts commenced. The tail sitter, nowknown as Flamingo was deemed the winning concept and was designed further during the detaileddesign phase of the DSE. Due to the fact that the Flamingo needs to perform the VATOL maneuver,many man-hours were spent on making a successful simulation of this procedure in order to verifythe stability and control of the aircraft.

The designed UAV meets all customer based requirements and most of the designer set require-ments. It is capable to operate in environments all over the world as well as being in contact withthe support system via a worldwide Internet connection. The Flamingo can land on areas of 10 by10 meters, using a vertical attitude take off and landing maneuver, deploy its payload and fly backto the base again, autonomously. Special care was taken to make the design as cost efficient and assustainable as possible. This was done by choosing a Wankel engine and minimizing the material andcomponents cost. Doing so makes the Flamingo significantly more fuel efficient and more sustainablethan the main competitor, the manned helicopter. The most critical structural components, suchas the the wing box, landing gear, engine mount, canard, cargo pod and the booms connecting thecanard to the main wing are designed in detail. Further, finite element modeling is used to designthe wing box and the booms in greater detail. The results from the aerodynamic model, using thevortex lattice method, and the flight dynamics simulation confirm that the aircraft is marginallydynamically unstable. However, the controller keeps it stable throughout the cruise and the durationof the VATOL maneuver.

One flamingo is capable of providing 300 beneficiaries with humanitarian aid during one mission.All beneficiaries now do have supplies for one week. A fleet of ten UAVs is able to provide 20000disaster stricken people with emergency aid during 420 missions in one week. It does this at a costof AC151 per mission. Table 8.1 lists some of the main characteristics of the Flamingo.

87

88 8 Conclusions

Aspect Value

Preliminary production Cost UAV 1 AC86000

Cost for one mission AC151

Fuel used for one mission 28.4kg

Number of missions per day 6

Number of crew required at airport 15

MTOW 7140N

Empty weight 4150N

Maximum payload (excl. cargo pod) 250kg, 1m3

Max altitude VATOL (250kg payload) 900m

Max altitude VATOL (100kg payload) 5200m

Range (with payload) 590km

Range (without payload) 940km

Maximum mission range 300km

Cruise altitude 3000m

Cruise speed 55m/s

Take off length 615m

Time for VATOL maneuver 5minTW -ratio 1.04(MTOW)/1.72(w.o. payload)

Span 7.45m

Total length 4.8m

Table 8.1: Flamingo characteristics

It can be concluded that the Flamingo UAV and the UAS of which it is a part, are feasible and couldform a useful tool in future humanitarian emergency relief. Especially because of its low productionand operation cost, it is a cheap alternative to the manned helicopter. It’s small size and payloadwill prove to be useful in reaching smaller communities in remote areas more cost efficiently. At thesame time being able to reach many different places in a short period of time thus providing a bettergeographical distribution of aid with respect to a conventional helicopter.The project objective statement of this group is:

Design a cheap and sustainable UAS to supply aid several times a day, to difficult to access disasterareas, in 11 weeks time within a group of 10 students.

The contents of this report clearly illustrate that this objective has been accomplished.

Chapter 9

Recommendations

During this project many aspects of the Flamingo have been designed and developed. However, dueto the time constraint of this project many more aspects still need to be looked in to. In this chaptera list recommendations for the further development of this UAV is given. First, the recommendationsregarding the organization of the project are given in section 9.1. Then, further recommendationsfor the actual design will be made. In section 9.2 the recommendations for the aerodynamic design,the aircraft stability and the performance are discussed. Section 9.3 the recommendations for thestructural design will be mentioned. In section 9.4 the recommendations for the simulation willbe discussed. The recommendations for the subsystem design are discussed in 9.5 and finally therecommendations for the operations and logistics can be found in section 9.6.

9.1 Recommendations for the project in general

Even though the project was well organized there is always room for improvement. During the finaldesign much work was inter depended, so communication about design progress was very important.Even though daily meetings were held, some unnecessary work was done because design changes arenot communicated through instantaneously. A better system for managing these variables and theirstatus, for example a notification board, is therefore recommended.

9.2 Recommendations for aerodynamics, stability and performance

Recommendations concerning aerodynamics are mainly aimed at higher levels of accuracy. TheVATOL maneuver is heavily dependent on the propeller induced air velocity. When introducing apropeller induced velocity field, which takes into account the non-homogeneous velocity field in thebody X-direction and the rotational component due to the propeller rotation; more accurate liftdistributions and flaperon/rudder deflection derivatives can be obtained. Moreover, the body canbe modeled using the actual body shape instead of a NACA0025 airfoil for better results. Due tothe TORNADO method, it is impossible to investigate the stall characteristics of the aircraft. It isadvised to use more advanced, analytical methods for further research incorporating the viscid natureof airflows. These methods can also enhance the drag estimation, especially that of the cargo pod.The effects of winglets are not investigated during the project. The winglets may turn out beneficialwhen the characteristics of the aerodynamic performance, structural design and cost are combined.Finally, the cargo pod will have a significant influence on the flight characteristics of the Flamingo.Further extensive research must be done in order to investigate the influence of the cargo pod andoptimize its shape.

Investigation of stability can also be enhanced using the aforementioned recommendations. In-troducing wind will also give better results concerning static hovering stability. The static stability

89

90 9 Recommendations

of the Flamingo during conventional flight must be investigated further in order to reach equilibriumwith smaller canard and flaperon deflections.

As always during aircraft design, the weight must be kept as low as possible. As, the proprotors area compromise between cruise and hovering flight; the maximum hovering altitude is limited. In orderto increase this maximum hovering altitude and enable worldwide emergency relief coverage withfull payload, the UAV weight must decrease. In this report, the mission performance is calculatedusing for example a constant value for the rate of climb and constant cruise altitude. More advanceddiscretized modeling of the mission will optimize the mission performance and increase the accuracyof the fuel consumption and total flight time estimation.

For the proprotor design, it is recommended to look deeper into the transitional phase. Now,cruise flight and hovering are considered separately and the transitional phase is interpolated. Toachieve higher accuracy, it is recommended to revise the aerodynamics of the proprotors during thetransitional phase. Blade Element Theories for fixed-wing aircraft and rotorcraft were used for theproprotor design. For more accuracy, it is recommended to look into other advanced theories ortheories with a different approach, such as the vortex theory.

9.3 Recommendations for structures

The structural parts of the Flamingo has already been designed in high levels of detail. However, thereare several things which have not been considered due to time constrains. Especially, the connectionsbetween different parts of the structure have not been looked in to in detail. It is assumed that theloads are transferred between the parts but the actual manner and the mechanisms involved havenot been looked in to. Furthermore, the structural strength is verified for only one part of the flightenvelope, while it should be checked for further critical scenarios dependent on both the load factorand velocity.

For the wing box in particular, the required strength to weight ratio of the top and the bottomplate was optimized using a numerical simulation, such that it resulted in the minimum possibleweight. This is achieved by applying ribs and stringers. The spars however have not been optimizedfor weight, but merely verified to ensure that they do not fail under the applied loads. Further iter-ations should be performed on all wing box computations to ensure the structure is not over sized.Moreover, the number of fasteners used for the spar caps have now been sized for buckling loads toensure no plastic deformation. The fasteners are also used to transfer the loads from the top andbottom plate to the spars, hence they should be sized for this as well.

The finite element models should be analyzed in much greater detail to include failure modes, suchas buckling and vibrational analysis. In this case, the use of PATRAN must be recommended overCATIA, as it allows for better mesh control and does not limit itself to solid finite element model-ing. Furthermore, PATRAN allows for more accurate load application. Especially, the finite elementmodel of the booms connecting the canard to the main wing should be re-analyzed in greater detail.Currently, they are treated as metal structures, where as each layer of carbon fibers should be modeledwith separate elements.

Further improvements to the finite element models computed during this project include applyinga temperature field and modeling its effects on the structural parts. Inter rivet buckling failure modeshould also be looked in both analytical and FEM computations. In FEM, it can be achieved byusing rivets to connect two solids in the CATIA model, rather than using surface contacts.

9.4 Recommendations for the simulation

The simulation that has been performed needs to be further developed as it will be the basis for theautopilot. Currently the model is based on the following assumptions:

9 Recommendations 91

• Flat, non-rotating earth

• No gyroscopic effects

• Steady atmosphere

• Sea level, standard atmosphere conditions

• Induced velocity is equal to mean induced velocity and is evenly distributed over main wing

These assumptions allowed to simplify the model while still allowing to verify the VATOL ma-neuver. This is not sufficient for the auto pilot. Of all these assumptions, only the flat, non-rotatingearth assumption will mostly be valid. All the other assumptions will prove to change the actual flightdynamics of the Flamingo. In the further development, these assumptions should either be validatedor need to be incorporated in the model. Especially the static atmosphere needs thorough modellingas, requirement D.5states that the Flamingo should be able to land in 7Bft wind conditions. Alsothe ISA atmospheric conditions should be considered, since the aircraft will also have to perform theVATOL maneuver at higher altitudes. The effect of the induced velocity field behind the propelleron the flight dynamics should be investigated aswell.

Perhaps the most important recommendation is to investigate the possibility to perform the tran-sition to vertical flight at constant altitude. The current pitch up maneuver brings the Flamingo toa relatively high altitude. This greatly increases the time in vertical flight. A constant altitude tran-sition to vertical attitude flight could greatly reduce the fuel consumption in the VATOL maneuver.

9.5 Recommendations for the subsystems

The subsystems have not been developed in detail in this project and require further design. Manycomponents are available off the shelf but their interaction has to be elaborated on. The fuel tanksshould be designed in further detail, especially as their layout is rather novel. An important detailthat should be further analyzed is how the fuel quantity should be measured in this novel bladdertype fuel tank. A proposed method is to include a system that can weigh the fuel tank, while takingthe load factor in to account. This information can then provide an estimate of the remaining fuel inthe tank. For the communication system, more research should be done in to the relay of informationbetween the UAVs. Currently, there are no details on how this system should work. The autopilotrequires thorough testing. Finally the sensor choice and position should be scrutinized. It is veryimportant that they give enough information to perform a controlled flight and landing.

9.6 Recommendations for operations and logistics

The operations and logistics are essential for the use of the UAS and for the success of the missionin terms of emergency relief. It is safe to say that lives depend on it. This report verifies that 10UAVs can fit within a A400M. However, the details of how this is done with respect to restrainingthe motion of the equipment during flight and the most efficient way of loading and unloading needsto be developed further. Also, the training of the personnel and how they will operate on the airportneeds to be looked into. It is currently assumed that the A400M delivers the UAS and that the actualgoods to be delivered to the disaster site will be transported in another aircraft or made availablethrough other means. The details of organizing this and how the goods are packed in the cargo boxeshas to be developed. Another aspect of the logistics that needs to be considered is how a group ofaircraft will fly with respect to each other. All UAVs should not land at the same time as the currentsupport system will not be able to handle this. It will be interesting to see if two Flamingos mightbe able to take off and land at the same time, reducing the pressure on a busy airport in emergencysituations.

92 9 Recommendations

Bibliography

AirSuppliers (15-06-2011), http://www.airsuppliers.com/Cessna_(Textron)_tires.htm.

Anderson, J. D. (2005), Introduction to Aerodynamics, 5 edn, McGraw-Hill.

Aplique Kit (16-06-2011), www.swri.org/4org/d10/its/ivs/pdfs/MARTIApplique.pdf.

BGAN Service Rate Pricing from Ground Control (16-06-2011), http://www.groundcontrol.com/BGAN_rate_plans.htm.

Bruhn, E. (1973), Analysis & Design of Flight Vehicle Structures, Jacobs publishing.

Calister, W. D. (2006), Materials Science and Engineering, Wiley.

CES Edupack Software (2011).

Currey, N. S. (1988), Aircraft Landing Gear Design: Principles and Practices, American Institute ofAeronautics and Astronautics.

D.E. Ellison, L. M. e. a. (1965), Usaf stability and control datcom, Technical report.

EAS (2003), CS-23 (Normal,Utility,Aerobatic,and Commuter Aeroplanes).

Ground Control (15-06-2011), http://www.groundcontrol.com/Explorer_727_Thrane_Thrane_

BGAN.htm.

Hardy Diesel Generators (16-06-2011), http://www.hardydiesel.com/

diesel-generators-35-115-kw.html.

M. Drela, H. Y. e. (2011), ‘Xflr5 v6.03’, http://xflr5.sourceforge.net/xflr5.htm.

Megson, T. (2007), Aircraft Structures for Engineering Students, Elsevier.

Melin, T. (2000), ‘Tornado, a vortex lattice matlab implementation for linear aerodynamic wingapplications’, Masters thesis.

Melkert, J. (2011), ‘Ae3200 design synthesis’, Aeropace Engineering BSc Study Guide.

MicroPilot MP2128 (14-06-2011), http://www.micropilot.com/products-mp21283x.htm.

Military Specification MIL-S-6073 (06-11-1945), http://www.westcoastpiet.com/images/

Construction/Military%20Specification%20MIL-S-6073_%20%20Aircraft%20Spruce.pdf.

93

94 BIBLIOGRAPHY

Mooij, E. (1994), Motion of a vehicle in a planetary atmosphere, Technical report, Delft Universityof Technology.

Mooij, E. (1997), Linear quadratic regulator design for an unpowered, winged re-entry vehicle, Tech-nical report, Delft University of Technology.

Narrative History of Operation Unified Response (2010), http://www.southcom.mil/appssc/

factFilesLarge.php?id=138.

Phillips, W. F. (2004), Mechanics of Flight, Wiley.

PrimeSense: Motion Control Beyond the Kinect (16-06-2011), http://blogs.pcmag.com/miller/2011/05/primesense_motion_control_beyo.php.

R44 Raven/Clipper Series Helicopters (17-05-2011), http://www.robinsonheli.com/rhc_r44_

raven_series.html.

R.C.Hibbeler (2005), Mechanics of Materials, Prentice Hall.

RIEGL LMS-Q160 (14-06-2011), http://www.riegl.com/products/airborne-scanning/

produktdetail/product/scanner/8/.

Robinson Model R-44 (11-05-2011), http://www.aviastar.org/helicopters_eng/robinson-44.

php.

Ruijgrok, G. (1996), Elements of Airplane Performance, Delft University press.

The rotary engine for 21st century general aviation G-200 (n.d.).

van Baten, T. and G.N.Saunders-Smits (January 2004), AE1-005STR-Aerospace Engineering Project,TU Delft, Faculty of Aerospace Engineering.

Appendices

95

Appendix A

Technical drawings

97

98 A Technical drawings

Figure A.1: Isometric view of the Flamingo

A Technical drawings 99

Figure A.2: Top view of the Flamingo

100 A Technical drawings

Figure A.3: Front view of the Flamingo

A Technical drawings 101

Figure A.4: Side view of the Flamingo without cargo pod

102 A Technical drawings

Figure A.5: Side view of the Flamingo with cargo pod

Appendix B

State Space System Matrix Inputs

A =

CXu 0 0 0 0 0 −CXe0 CXex −CXey CXez0 CYv 0 0 0 0 CYe0 CYex CYey CYez0 0 CZw 0 0 0 CZe0 −CZe0 −CZe0 CZe00 0 0 Cpp Cpq Cpr 0 0 0 00 0 0 Cqp Cqq Cqr 0 0 0 00 0 0 Crp Crq Crr 0 0 0 00 0 0 −Ce0p −Ce0q −Ce0r 0 0 0 00 0 0 −Cexp −Cexq −Cexr 0 0 0 00 0 0 −Ceyp −Ceyq −Ceyr 0 0 0 00 0 0 −Cezp −Cezq −Cezr 0 0 0 0

(B.1)

CXu = 2CLααtan(γ)qdynSmV

CXe0 = 2gey0

CXex = 2gez0CXey = 2ge00

CXez = 2gex0

(B.2)

CYv = −2CYβqdynSmV

CYe0 = 2gex0

CYex = 2ge00

CYey = 2gez0CYez = 2gey0

(B.3)

CZu = −2CLααqdynSmV

CZw = CLαqdynSmV

CZe0 = 2ge00

CZex = 2gex0

CZey = 2gey0

CZez = 2gez0

(B.4)

Cpp = CLpqdynIzz

IzzIxx−I2zx

Sb2 + Cnpqdyn

IxzIzzIxx−I2

zx

Cpr = ClrIzz

IzzIxx−I2zx

Sb2

Cqq = Cmq1IyyqdynSc

Crp = CnpqdynIxx

IzzIxx−I2zx

Sb2

Crr = CmrqdynIxx

IzzIxx−I2zx

Sb2 + Clrqdyn

IxzIzzIxx−I2

zx

Sb2

(B.5)

103

104 B State Space System Matrix Inputs

Ce0p = 12ex0

Ce0q = 12ey0

Ce0r = 12ez0

Cexp = −12e00

Cexq = 12ez0

Cexr = −12ey0

Ceyp = −12ez0

Ceyq = −12e00

Ceyr = 12ex0

Cezp = 12ey0

Cezq = −12ex0

Cezr = −12e00

(B.6)

B =

0 0 0 0 CXδT 00 0 0 CYδr 0 0

CZδc 0 CZδe 0 0 00 Clδa 0 0 0 0

Cmδc 0 Cmδe 0 0 00 0 0 Cnδr 0 Cnδ∆T0 0 0 0 0 00 0 0 0 0 00 0 0 0 0 00 0 0 0 0 0

(B.7)

The values in the B matrix are computed using the aerodynamic model, discussed in chapter 3.2except for the engine coefficients, those can be found in equation B.8. Note that δT and δ∆T takevalues between ±1 so these coefficients are multiplied by the maximum thrust.

CXδT = Tmaxm

Cnδ∆T =Tmaxlyeng

Izz

(B.8)

Appendix C

Compliance Matrix

ID Requirement Value Com1 Justification

A.1 Non toxic materi-als

- Yes The materials used for structures are all nontoxic as explained in chapter 4.

A.2 Recycling - No Except for the cargo pod and booms all otherstructures are made of a material that can berecycled as explained in chapter 4.

B.1 Payload delivery - Yes The UAV is able to carry payload and deployit autonomously.

B.1.1 Deploymentmechanism

- Yes The deployment mechanism is explained in sec-tion 4.4.

B.1.2 Guidance system - Yes Information on guidance system is explained insection 5.4.

B.1.2.1 Navigation sys-tem

- Yes Navigation system is explained in detail in sec-tion 6.3.

B.1.3 Report status - Yes This is explained in detail in section 6.4.

B.1.3.1 Communicationsystem

- Yes The detailed information on communicationsystem is explained in section 6.4.

B.1.4 Power supply - Yes The explanation of power supply is explainedin section 6.1.

B.2 Provide visuals - Yes The UAV is capable of taking pictures to de-cide a suitable spot to land as explained in sec-tion 6.5.

B.2.1 Remote sensingequipment

- Yes The UAV is equipped with several sensors asexplained in section 6.3.

B.3 Mission safety - Yes The UAV can perform mission safely.

B.3.1 Safety of people - N.A. Ensuring the safety of the people on the land-ing site should be further considered 2.4.

B.3.2 Safety of system - Yes The system stays safe during the UAV opera-tion.

B.4 Land on roughsurface

- Yes The pod can land on a reasonably rough sur-face.

C.1 Maximum Pay-load mass

≤ 250 kg Yes The structures have been analysed such thatthey can handle loads when carrying a payloadof 250kg as explained in chapter 4.

1Compliance: Yes, No or Not Applicable (N.A.)

105

106 C Compliance Matrix

C.2 Maximum Pay-load volume

≤ 1 m3 Yes The pod is designed such that it can store 1m3

volume as explained in section 4.3.

C.3 Land area size 10 m by10 m

Yes The UAV can land in 10m by 10m area.

C.3.1 Maximum Span <14.14 m Yes The span of the wing is 7.454m which is quitebelow the maximum span allowed.

C.4 Transport UASand supportsystem

10 UAVs+ SS 2 inA400M

Yes An A400M can transport 10 UAV units, sup-port equipment and staff to disaster site as ex-plained in chapter 7.

C.4.1 Total UAS vol-ume

340 m3 Yes The total volume of all 10 UAVs componentsis less than 340 m3. meaning it can fit in aA400m

D.1 Range at maxi-mum payload

>400 km Yes The UAV is able to travel 200 km with payloadto disaster site and then travel another 200kmback to the airport without payload.

D.2 Sorties per day >4 Yes Each UAV can perform 6 sorties a day.

D.2.1 Cruise speed >32 m/s

Yes The cruise speed of the UAV is 55m/s.

D.3 Take off length <1000 m Yes The takeoff length is 615m as computed in sec-tion 3.6.7.

D.4 Maximum pay-load acceleration

<2 g Yes The load factor is below 2g as explained inchapter 5.

D.4.1 Power systems - Yes The engine used is a rotary engine and no rock-ets or boosters are used as explained in section6.1.

D.5 Wind speed >7 Bft Yes All analysis was performed taking into accountwinds of 7 Beaufort.

D.6 VTOL - Yes The UAV can perform the vertical takeoff andlanding which is explained in detail in chapter5

D.7 Ground slope ≤ 10◦ Yes The pod is capable of landing on a slope of 10◦

without tipping over as explained in section 4.3

D.8 Landing duringnight

- Yes The UAV can land during night as explainedin section 6.3.

D.8.1 Sensor system - Yes The UAV is equipped with appropriate sensorsto land at night as explained in section 6.3.

D.9 Cost per sortie <AC500 Yes This is explained in section 3.7.

D.10 Unit productioncost

<AC250,000 Yes It is explained in chapter 7

D.11 World coverage - Yes The UAV can remain functional in very differ-ent environments.

D.11.1 Service ceiling 6000 m No The maximum altitude for the UAV to hoveris 5200m with a payload of 100kg so it is notpossible to reach the highest settlements in theworld as explained in section 3.6.

D.11.2 Maximum alti-tude

9000 m Yes The maximum altitude to in cruise with pay-load is 17500m as explained in section 3.6.

2Support System

C Compliance Matrix 107

D.11.3 Temperaturerange

-60◦C to60◦C

N.A. The material for each structure has been se-lected such that it can operate in an as widetemperature range as possible, but especiallthe electrical systems need to be varified atthese temperatures.

D.11.4 Maximum humid-ity

80% N.A. This has not been looked into detail due totime constraints.

D.11.5 Sand tolerance3 1 g/m3 N.A. The material has not been checked for sandtolerance due to time constraints.

D.11.1.1 Power system - Yes The rotary engine can operate without compli-cations under stated conditions as explained insection 6.1.

D.11.1.3 Material choice - Yes The materials used for the UAV can functionin all possible environments.

E.1 CO2 emission re-duction

>75 % Yes The 75 % CO2 reduction is explained in section2.6.

E.1.1 Maximum fuelmass per sortie

29.4 kg Yes The fuel mass used per sortie is 28.4kg whichmeans the UAV meats up with the CO2 reduc-tion as explained in section 3.7.

F.1 Certification4 - N.A. Except for the landing gear it has not beenlooked in detail due to time constraints.

G.1 Manpower 10 people Yes The design and analysis is a collaborative effortof 10 people.

Table C.1: Compliance Matrix

3ISO 12103-1 A4/A24CS-23 / FAR-23

108 C Compliance Matrix