Landing Gear Design for Blended Wing Body Flight Test ...

Landing Gear Design for Blended Wing Body Flight TestDemonstrators

Fabrizio Rizzi

Thesis to obtain the Master of Science Degree in

Aerospace Engineering

Supervisor: Prof. Afzal Suleman

Examination Committee

Chairperson: Prof. Fernando José Parracho LauSupervisor: Prof. Afzal Suleman

Member of the Committee: Dr. Frederico José Prata Rente Reis Afonso

November 2018

PRIVATE AND CONFIDENTIAL©Bombardier Inc. or its subsidiaries. All right reserved.

Abstract

The Blended Wing Body (BWB) configuration is a hybrid shape with unique features capable to takebenefits from both flying wing and conventional aircraft. The poof of an unconventional concept confi-guration requires flight tests and validation on small Flight Test Demonstrators, avoiding cost and riskrelated to the use of a full scale model.

A first 7% scaled Unmanned Aerial Vehicle (UAV) for a new generation BWB aircraft, has beendesigned and flight tested at the Center for Aerospace Research (CfAR), operating on unprepared grassrunways (GRS). The need to evaluate the UAV, fully controlled with an autopilot, even in the criticalphases of take-off and landing on concrete runways (CON), has required the design and integration of anundercarriage system.

One of the main goals of the present thesis work is to carry out the design and development of alanding gear for FTV7%, including the integration of the system into the aircraft, ground testing andmock-ups preparation for a wheeled flight test campaign.

In the process of scaling towards the faithfully representation of all the aspect that concern a fullscale aircraft, the design of the landing gear has been started in parallel, even for a new 16.5% scale,involving some design similarities and the sizing of additional mechanical subsystems, such as dedicatedsuspension and braking systems.

At last, the basic take-off and landing performance evaluation of the landing gears, designed for bothFTV7% and FTV16.5%, is presented highlighting the influence of some design parameters.

Keywords: BWB, Undercarriages, Design, Scaling, Performance evaluation.

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Resumo

A Blended Wing Body (BWB) é uma configuração com forma híbrida e com características únicas quelhe permitem beneficiar das vantagens da configuração convencional e em asa voadora. A prova doconceito de uma configuração não convencional requer ensaios em voo e validação através de Flight TestDemonstrators, evitando assim custos adicionais e riscos associados a ensaios à escala real.

Um primeiro modelo Unmanned Aerial Vehicle (UAV) à escala de 7% de uma aeronave com confi-guração BWB tem sido desenvolvido e submetido a ensaios de voo pelo Centre for Aerospace Research(CfAR), os quais têm sido feitos em pistas de aterragem indevidamente preparadas. A necessidade deavaliar um UAV controlado com um piloto automático em cenários críticos de descolagem e aterragemem pistas cimentadas requer o projecto e integração de um trem de aterragem.

Um dos objectivos principais deste trabalho consiste no projecto do trem de aterragem do FTV7%,incluindo a integração do sistema na aeronave, ensaios no solo e a preparação duma maquete para acampanha de ensaios de voo.

Paralelamente, durante o processo de redução de escala, foi desenvolvido o projecto do trem deaterragem de um modelo à escala de 16.5%. Este contem semelhanças com anterior mas com subsistemasmecânicos adicionais, tais como uma suspensão e um sistema de travagem.

Por fim, o desempenho do trem de aterragem durante a descolagem e aterragem são avaliados e pro-jectados para o FTV7% e FTV16.5%, são apresentados com foco na influência de alguns parâmetros deprojecto.

Palavras-chave: BWB, trem de aterragem, projecto, modelos à escala, avaliação de desempenho.

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List of Figures

2.1 Landing gear for different aircraft applications. [6] . . . . . . . . . . . . . . . . . . . . . . 52.2 Ground clearance requirement during take-off rotation. [10] . . . . . . . . . . . . . . . . . 72.3 Take-off rotation and tip-back parameters. [10] . . . . . . . . . . . . . . . . . . . . . . . . 72.4 Landing gear application for several UAVs . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.1 Interface of the StatData feature of the Flight Test Data Processing software . . . . . . . 123.2 CAD assembly of the designed main landing gear for FTV7 . . . . . . . . . . . . . . . . . 143.3 Option 1: Skymaster wheels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153.4 Option 2: Turnigy HK wheels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153.5 SW structural simulations of the Al2024-1/8" leaf strut . . . . . . . . . . . . . . . . . . . 163.6 Drop Test of the Al2024 1/8" performed at CfAR . . . . . . . . . . . . . . . . . . . . . . . 163.7 Geometric parameters for the leaf strut design . . . . . . . . . . . . . . . . . . . . . . . . 163.8 Design taper parameters for the leaf strut. [8] . . . . . . . . . . . . . . . . . . . . . . . . . 173.9 Integration of the main gear with the mid bay of the aircraft . . . . . . . . . . . . . . . . 173.10 Crack on the lower bending radius of the leaf strut, after the first bending machine process 183.11 Final bending hot process of the main gear struts . . . . . . . . . . . . . . . . . . . . . . . 183.12 Heat treatment of the leaf strut. [27] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 183.13 Displacement of the main gar strut with a static load corresponding to the MTOW of FTV 7 193.14 Test rig with no loading applied . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203.15 Test rig with static loading corresponding to the MTOW of the aircraft . . . . . . . . . . 203.16 Load displacement curve for the main gear leaf strut . . . . . . . . . . . . . . . . . . . . . 213.17 CAD assembly of the designed nose gear for FTV7 . . . . . . . . . . . . . . . . . . . . . . 223.18 Nose strut with static load corresponding to the aircraft MTOW . . . . . . . . . . . . . . 233.19 Nose strut with bottoming force, corresponding to the impact in hard landings . . . . . . 233.20 CAD design of the custom components for the nose gear of FTV7 . . . . . . . . . . . . . . 243.21 Geometry and kinematics of the steering system . . . . . . . . . . . . . . . . . . . . . . . 253.22 Turning radius for the steering system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 253.23 Steering servo components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 263.24 Mechanical steering system components . . . . . . . . . . . . . . . . . . . . . . . . . . . . 263.25 Static test for the characterization of the nose gear strut: the test rig is represented on the

left, the resulting load/deformation curve on the right. . . . . . . . . . . . . . . . . . . . . 273.26 External and internal integration of the main landing gear with the airframe . . . . . . . 283.27 Nose landing gear external and internal integration with the airframe . . . . . . . . . . . . 293.28 Weight and Cost distributions for the production of one landing gear set for FTV7 . . . . 30

4.1 Tuning of the steering system control parameters for the autopilot . . . . . . . . . . . . . 324.2 Component list and mounting drawing for the Nose Gear strut assembly 1 . . . . . . . . . 334.3 Landing overview using the Flight test data processing software . . . . . . . . . . . . . . . 35

5.1 Work break-down structure of the landing gear design . . . . . . . . . . . . . . . . . . . . 385.2 Multidisciplinary process of the landing gear design . . . . . . . . . . . . . . . . . . . . . . 395.3 Basic requirements for the 16.5% Landing Gear design . . . . . . . . . . . . . . . . . . . . 405.4 Equilibrium about the main gear contact point at take-off rotation . . . . . . . . . . . . . 415.5 Sketch of the swivel component, necessary to rotate the main gear position . . . . . . . . 425.6 Take-off rotation ground clearance requirement . . . . . . . . . . . . . . . . . . . . . . . . 435.7 Geometric parameters involved in the determination of the turnover angle . . . . . . . . . 43

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5.8 Design procedures for impact loading condition: on the left the scaling design procedurewith influence of the parachute rate of descent; on the right the classical design procedurewith influence of the design wheel travel on landing . . . . . . . . . . . . . . . . . . . . . . 47

5.9 Sketches of the concepts for the internal layout configuration of the main landing gear. . . 485.10 Off-the-shelf tires from the UAV and ultralight aircraft market . . . . . . . . . . . . . . . 495.11 Off-the-shelf wheels from the UAV and ultralight aircraft market . . . . . . . . . . . . . . 505.12 Off-the-shelf brake solutions for FTV16.5 applications . . . . . . . . . . . . . . . . . . . . 505.13 Off-the-shelf shock absorbers for FTV16.5 applications . . . . . . . . . . . . . . . . . . . . 515.14 Conceptual geometry of the main gear strut leg . . . . . . . . . . . . . . . . . . . . . . . . 515.15 Initial structural evaluation of different custom main gear struts . . . . . . . . . . . . . . 525.16 Conceptual sketches of the carbon fiber strut . . . . . . . . . . . . . . . . . . . . . . . . . 525.17 Sketches of the concepts for the nose landing gear for FTV16.5 . . . . . . . . . . . . . . . 535.18 Wheel assembly solutions for the nose gear of FTV16.5 . . . . . . . . . . . . . . . . . . . . 545.19 Concepts for the steering system of the landing gear for FTV16.5: Rack and Pinion me-

chanical transmission on the left and Pulley-Belt mechanism on the right . . . . . . . . . 545.20 Test rig for shock absorbers: it shows all the components needed for the test rig assembly. 555.21 Test rig concepts: on the left the Speed Rating Test rig, on the right the Drop Test rig. . . 565.22 Feature importance for the Main gear concept selection . . . . . . . . . . . . . . . . . . . 575.23 Feature importance for the Nose gear concept selection . . . . . . . . . . . . . . . . . . . . 575.24 Concepts selected for the main gear (on the left) and for the nose gear (on the right) . . 585.25 Procurement of the main gear wheel assembly . . . . . . . . . . . . . . . . . . . . . . . . . 595.26 Procurement of the nose gear wheel assembly . . . . . . . . . . . . . . . . . . . . . . . . . 595.27 Off-the-shelf shock absorber for the nose and main landing gear . . . . . . . . . . . . . . . 605.28 Internal chambers of the shock absorber [38] . . . . . . . . . . . . . . . . . . . . . . . . . . 605.29 Shock load for the FOX 3 DPS 6.5"×1.5", selected for the Main gear application . . . . . 615.30 Shock load for the FOX 3 DPS 5.5"×1", selected for the Nose gear application . . . . . . . 615.31 Design space for the main landing gear inside the airframe of FTV16.5 . . . . . . . . . . . 625.32 Preliminary design geometry and load distribution of the main gear selected concept. . . . 625.33 Braking system for ultralight applications: internal caliper plus brake disk. [43] . . . . . . 645.34 Braking system from bike applications applied to the X-48 UAV. [17] . . . . . . . . . . . . 655.35 Design space for the nose landing gear inside the airframe of FTV16.5 . . . . . . . . . . . 655.36 Preliminary design geometry and load distribution for the nose gear selected concept. . . 665.37 Pulley-belt steering system design parameters. [37] . . . . . . . . . . . . . . . . . . . . . . 665.38 Weight and Cost estimation for the production of one landing gear set for FTV16.5 . . . . 68

6.1 Aircraft taxiing operation over a 1-cosine modeled runway . . . . . . . . . . . . . . . . . . 706.2 Vertical acceleration during taxiing over a 1-cos dip runway for the FTV7 and FTV16.5 . 716.3 Heave response during taxiing over a 1-cos dip runway for the FTV7 and FTV16.5 . . . . 716.4 Aircraft model during landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 726.5 Vertical impact deceleration on the main gear of FTV7 and FTV16.5 . . . . . . . . . . . . 736.6 Vertical displacements on landing on the main landing gear of FTV7 and FTV16.5 . . . . 73

B.1 Main Gear assembly drawing 1: leaf strut and attachment with the airframe . . . . . . . . 86B.2 Main Gear assembly drawing 2: wheel assembly connection with the leaf strut . . . . . . . 86B.3 Nose gear assembly drawing 2: wheel assembly and piston strut . . . . . . . . . . . . . . . 87B.4 Nose gear assembly drawing 3: steering system . . . . . . . . . . . . . . . . . . . . . . . . 87B.5 Shock absorber Test Rig assembly drawing 1 . . . . . . . . . . . . . . . . . . . . . . . . . 88B.6 Shock absorber Test Rig assembly drawing 2 . . . . . . . . . . . . . . . . . . . . . . . . . 88

C.1 Rigid aircraft with equivalent model of undercarriages during taxiing over the designrunway.[39] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

C.2 Simulink model for the heave response on time domain during landing . . . . . . . . . . . 91

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List of Tables

2.1 Pros and Cons of the most used landing gear types [10] . . . . . . . . . . . . . . . . . . . 6

3.1 General requirements for the landing gear design for FTV7 . . . . . . . . . . . . . . . . . 123.2 Aircraft and performance requirement for FTV7 . . . . . . . . . . . . . . . . . . . . . . . 133.3 Geometry requirements for FTV7 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133.4 Specific requirements for the main and nose gear of FTV7 . . . . . . . . . . . . . . . . . . 143.5 Analytical Hierarchical table for tire and wheel selection of FTV7 main landing gear. . . . 153.6 Geometric parameter determination of the custom leaf strut design . . . . . . . . . . . . . 173.7 Resulting displacements from Static Tests with increasing weight: (0 : 3 : 30kg) . . . . . . 203.8 Linearized stiffness from experimental, analytical and computational tests . . . . . . . . . 213.9 Drop test setting and results for different simulated landing cases . . . . . . . . . . . . . . 223.10 Linearized stiffness coefficients for the nose strut during soft and hard performances . . . 273.11 Main landing gear position with respect to the Fuselage Station . . . . . . . . . . . . . . . 28

5.1 Set of landing gear configuration for different position of the c.g. . . . . . . . . . . . . . . 425.2 Track and turnover angle for each main gear configuration . . . . . . . . . . . . . . . . . . 455.3 Static loading cases on the nose and main gear for each landing gear configuration . . . . 455.4 Total loading cases on the nose and main gear for each landing gear configuration . . . . . 455.5 Scaling design procedure: impact accelerations developed on nominal and parachute landings 465.6 Classic design procedure: impact accelerations developed on nominal and parachute landings 465.7 Analytical hierarchical process table for the Main Gear concept selection . . . . . . . . . . 575.8 Analytical hierarchical process table for the Nose Gear concept selection . . . . . . . . . . 585.9 Braking distance calculation according to FAR 23 regulations . . . . . . . . . . . . . . . . 635.10 Calculation of the parameters for the preliminary sizing of the braking system . . . . . . . 645.11 Preliminary selection of the pulley-belt steering system . . . . . . . . . . . . . . . . . . . . 67

6.1 Input for the taxiing dynamic model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

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Contents

List of Figures v

List of Tables viii

1 Introduction 11.1 Background and Motivations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2 Scope of the Thesis Work . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.3 Contributions at the Center for Aerospace Research . . . . . . . . . . . . . . . . . . . . . 21.4 Collaborations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.5 Layout of the Thesis document . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2 Guidelines for Landing Gear Design 52.1 Design keys and variables for the Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . 52.2 Regulations for Landing Gear Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.2.1 Design limit parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82.2.2 Testing procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.3 State of the art of Landing Gear for similar aircraft applications . . . . . . . . . . . . . . 102.3.1 References for FTV7% . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102.3.2 References for FTV16.5% . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3 Landing Gear Design and Development for FTV7% 113.1 Requirement list . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3.1.1 General requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.1.2 Aircraft and Performance requirements . . . . . . . . . . . . . . . . . . . . . . . . 123.1.3 Geometry requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133.1.4 Main and Nose gear requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.2 Design and development of the main gear assembly . . . . . . . . . . . . . . . . . . . . . . 143.2.1 Main gear features . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143.2.2 Wheels and Tires for the main gear . . . . . . . . . . . . . . . . . . . . . . . . . . 153.2.3 Design of the main gear strut . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 153.2.4 Testing of the main gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

3.3 Design and development of the nose gear assembly . . . . . . . . . . . . . . . . . . . . . . 223.3.1 Nose gear features . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 223.3.2 Component off-the-shelf for the nose gear . . . . . . . . . . . . . . . . . . . . . . . 233.3.3 Design of additional custom components . . . . . . . . . . . . . . . . . . . . . . . . 233.3.4 Design of the steering system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 243.3.5 Testing of the nose gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

3.4 Integration of the landing gear with FTV7 . . . . . . . . . . . . . . . . . . . . . . . . . . . 273.4.1 Integration of the main gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 273.4.2 Integration of the nose gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 283.4.3 Belly Pan design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

3.5 Weight, cost and conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

4 Ground mock-ups and flight test planning for FTV7% 314.1 Preparation of the aircraft for ground and flight testing . . . . . . . . . . . . . . . . . . . 31

4.1.1 Ground mock-ups . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 314.1.2 Landing gear toolkit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

4.2 Ground testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

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4.2.1 Taxiing of the FTVs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334.2.2 Take off run testing of FTV 2B . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.2.3 Improvements on the landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

4.3 Flight test planning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 344.3.1 Flight test data processing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

5 Landing Gear design for FTV16.5% 375.1 Design process for the 16.5% Landing Gear system . . . . . . . . . . . . . . . . . . . . . . 375.2 Basic features of the 16.5% Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

5.2.1 Landing Gear position and aircraft center of gravity . . . . . . . . . . . . . . . . . 405.2.2 Landing Gear height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 425.2.3 Track and Turnover angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 435.2.4 Design loading conditions for the nose and main gear . . . . . . . . . . . . . . . . . 455.2.5 Impact loading condition for the landing gear . . . . . . . . . . . . . . . . . . . . . 45

5.3 Conceptual Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 475.3.1 Main gear concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 475.3.2 Nose gear concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 535.3.3 Test rig concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55

5.4 Preliminary Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 565.4.1 Selection of the concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 575.4.2 Procurement and testing procedures for the off-the-shelf components . . . . . . . . 595.4.3 Preliminary design of the main gear . . . . . . . . . . . . . . . . . . . . . . . . . . 625.4.4 Preliminary design of the nose gear . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

5.5 Weight, cost estimation and conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

6 Basic performance evaluation of aircrafts with a landing gear system 696.1 Response of the aircraft with landing gear during the typical ground maneuvers . . . . . . 69

6.1.1 Taxiing dynamic model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 696.1.2 Taxiing performance evaluation and comparison for FTV7 and FTV16.5 . . . . . . 70

6.2 Response of the aircraft with landing gear during a typical design landing . . . . . . . . . 726.2.1 2-points landing dynamic model . . . . . . . . . . . . . . . . . . . . . . . . . . . . 726.2.2 Landing performance evaluation and comparison for FTV7 and FTV16.5 . . . . . 73

7 Conclusions and Future developments 757.1 Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 757.2 Future developments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

A Landing gear Matlab parameter calculator 77A.1 Input parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77A.2 Output parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

B CAD assembly drawings 85B.1 Assembly drawings and Bill of components for the Landing Gear for FTV7% . . . . . . . 85B.2 Assembly drawings for the Shock Absorber Test Rig . . . . . . . . . . . . . . . . . . . . . 85

C Performance evaluation 89C.1 Taxiing Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89C.2 Landing simulink model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91

Bibliography 93

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Chapter 1

Introduction

1.1 Background and Motivations

The design and development of Unmanned Aerial Vehicles (UAVs) is the first step towards the proof offeasibility of new generation aircrafts. Instead of building an expensive full scale flight test demonstra-tor, the general approach is to scale the model and demonstrate that the configuration and associatedtechnologies warrant the development of a full-scale, certifiable aircraft. Experimental data for scaledmodels are used to define and review the basic characteristics of full-scale aircrafts, verify theoreticallypredicted behaviour and provide support for making decisions in low time, cost and risk. [1]

In this framework, the Canadian Aircraft Companies Bombardier and Quaternion Aerospace aremoving towards the development of an advanced, unconventional, BlendedWing Body (BWB) business jetwith an estimated entry into commercial service in 2035. [2] The scaling process is a step-by-step progressthat requires the development of small flight test vehicles (FTVs) in different scales and configurations,in order to test the effectiveness of each system and the affected behaviour of the aircraft during flight.The first scaled representation of the new generation aircraft was built and flight tested since 2016 atCenter of Aerospace Research (CfAR), with the collaboration of the University of Victoria (UvIC) inBritish Columbia - Canada. It is a 7% of the full scale aircraft, designed and developed having in mindthat the scaling of the physics of such a complex system goes far beyond merely scaling down the size.[3]

The objective of the first configurations of FTV 7% was to collect aerodynamic data from flight testsin order to validate and improve the control laws, used from the autopilot to maneuver the aircraft. Atthis time the phases of take-off and landing were not important to evaluate: the aircraft was catapultedinto the air using a shoot launcher and the landing was performed using the belly pan of the aircraft.

The next generation has required a lower risk to damage the vital components of the aircraft in termsof airframe and systems and the necessity to test the aircraft’s behaviour and control even during thecritical phases of take-off and landing. In order to accomplish these new needs towards an improvesfaithfullness of the represented scaled aircraft, a landing gear system has been needed and the groundoperation of the aircraft has moved from Grass airstrip to paved runways. 1

In parallel with the flight tests and demonstrations of expected performance for a wheeled configura-tion of FTV 7%, the Center for Aerospace research has been working on the design and development ofa larger scale flight test demonstrator that represents the 16.5% of the full aircraft size. The new scalerequires take-off and landing on paved concrete runways, so the design of a detailed landing gear systemis one of the hard-points that conditions the final layout and behaviour of the aircraft. The design andsizing of a landing gear system for such a large scale of UAV will provide an additional value towardsthe development of the new generation regional jet aircraft, now enabled to handle and dissipate impactshock forces with an adequate suspension system, to be stably maneuvered on the ground thanks to areliable steering system and to be stopped in a specific distance by using a certified braking system. [4]

1The aviation code for a grass runway is GRS,for a paved concrete runway is CON.

1

Fabrizio Rizzi: Landing Gear Design for Blended Wing Body Flight Test Demonstrators

1.2 Scope of the Thesis WorkThe main purpose of the present thesis work is to carry out the design and development of both under-carriages for two different scales of UAVs, that represent the new blended wing body business jet. For thesmaller flight test vehicle application, it is required that the landing gear enable the aircraft to be testedin a wheeled configuration with steering and braking capabilities linked to the operational conditions ofthe selected airstrip for testing. The landing gear has been designed, manufactured and integrated withthe airframe, including many aspects of aircraft design, fabrication, testing and analysis.

The landing gear for the larger scale demonstrator, instead, follows the general time-line and scheduleof the other design teams of the aircraft and a constant interchange of data and information is necessaryfor the development of a feasible aircraft concept with all the required systems and subsystems statedin the flying demo spec document [2]. The undercarriage system for FTV 16.5% is demanded to havea dedicated suspension in order to manage and dissipate the impact loads developed on take-off andlanding, and a detailed braking system to stop the aircraft within a certain distance.

The project also focuses on the design and development of custom test rigs in order to test theproperties of the components used, as well as prove the integrity and functionality of the designed landinggear in both static and impact load case scenarios.

All the phases of the work have been accompanied with the necessary documentation containing all thetechnical information of the landing gear, including component list definition, CAD drawings, updatedrequirement spreadsheets and support material for simulations and testing.

At the end of the present thesis work, the reader should be aware of all the unique challenges thatthe design of a landing gear offers, due to the multidisciplinary nature of the design process and thenecessity to evaluate structural behaviours and general performances since the first phases of preliminarydesign. He will be even conscious of all the ground mock-ups operations required after the installation ofa landing gear, that enable the wheeled aircraft to be fully controlled, in all the phases of flight, by anautopilot.

1.3 Contributions at the Center for Aerospace ResearchThe contributions resulting from the present thesis work are attributable to the area of design, integrationand performance evaluation of unmanned air vehicles equipped with landing gear. In particular theCenter for Aerospace Research is now provided of the following main outcomes, that are the result ofseven months of design work performed from March until September 2018:

• Design Excel spreadsheet: It contains all the critical requirements and key variables, necessaryfor the whole design process of a landing gear system.

• MatLab® Sizing tool: The software receives some input parameters stated in the Excel spreads-heet, and automatically calculates the resulting preliminary valid geometry that respects all therequirement verifications. The tool includes also sections to execute preliminary calculations forthe mechanical subsystems required to a landing gear.

• MatLab® & Simulink performance evaluation tool: The software can be used as a valuabletool for the heave performance evaluation of the landing gear during take-off and landing since thebeginning of the preliminary design and so, it can guide and demonstrate the feasibility of somedesign decisions.

• Checklist and hardware toolkit for the landing gear of FTV7%: All the landing gear faste-ners and components, designed and manufactured for FTV7%, have been organized in a dedicatedtoolkit box, with part number descriptions and all the necessary mounting drawings. A checklistfor the landing gear spare parts has been prepared as well, in order to facilitate the operations ofpre-flight checking and repairing of damaged components.

1.4 CollaborationsAll the design and realization phases of the project have been supported by the collaboration withseveral entities. First of all, the Center for Aerospace Research, located in an hangar of the InternationalVictoria Airport in Sidney and built in 2012 by the Professor Afzal Suleman in collaboration with theUniversity of Victoria, has represented the physical location where all the design process and ground

2

1 – Introduction

testing phases have been performed. The Center is specialized for UAV design in Western Canada andits shop mechanical machines have been extremely helpful to characterize the experimental nature ofdesign systems for Unmanned Aerial Vehicles. The Center’s human resources are organized in severalteams and a close collaboration has been needed with each of them that affects the landing gear design,such as the Airframe design, Recovery system and Control system teams. Ground mock-ups and flighttest preparations for the wheeled configuration of FTV 7% have been performed in collaboration withthe CfAR Flight Test Crew composed by Stephen Warwick, Jenner Richards, Sean Bazzocchi and MaxRukosuyev.

A constant reporting of the major design outcomes and progress on the production and manufactureside of the project, has been presented to the Bombardier stakeholders. All the activity for the landinggear project have been resumed in weekly progress reports, support technical documentation (quarterlyreports and coordination memos) and meetings every two weeks in order to verify and double check thefeasibility of the designed landing gear, ensuring that the design decisions made were not drasticallychanging the desired behaviour of the full scale aircraft.

The manufacturing phase of the landing gear for FTV 7% has been supported by several local com-panies that have provided the necessary mechanical instrumentation. The production of the leaf strutfor the main landing gear has been supported by the facilities of the companies Stark CNC and WesternEdison, specialized for waterjet cutting and bending metal components, and Pyrotek Aerospace, specia-lized for the heat processes of metal parts for aircraft applications. The manufacturing of needed designcomponents for the nose landing gear, as well as modifications for off-the-shelf components, have beenachieved using the manual lathe and mill machines provided at the Mechanical Laboratory of Universityof Victoria.

The design of the landing gear for FTV 16.5% has constantly been characterized by back and for-ward exchange of landing gear information with the major companies identified as suppliers for feasibleoff-the-shelf components, such as Matco Mfg and Aircraft Spruce Canada specialized for wheel assemblycomponents and Marc-Ingegno Italy ,Vorsprung and Fox, leaders of different suspension system applica-tions.

1.5 Layout of the Thesis documentThe organization of the thesis work in the present document tries to recall the chronology and evolutionof the design process for landing gear applications from small UAVs to larger scale demonstrators. Insome cases the work has been performed in parallel and so necessary rearrangements have been done inorder to preserve the reading flow. The resulting structure is described in the following itemize:

• Chapter 2: The general guidelines for the landing gear design for UAV applications, used for both7% and 16.5% FTVs, are presented. The chapter begins with an initial description of the designkeys and variables and specifications used to limit some design parameters. Then a state of the artof landing gear design applications, currently used for aircraft comparable with the two flight testdemonstrators, is described.

• Chapter 3: The third chapter shows all the design and development processes performed for thelanding gear of FTV7%. The main important requirements, that have conditioned all the design,are described and lead to the definition of the preliminary feasible layout of the landing gear. Thenthe design and production phases for both main and nose gear are explained in detail, includingthe required integration phase with the airframe of two FTV7% provided at the Center.

• Chapter 4: This chapter describes all the ground mock-ups and ground testing necessary tovalidate the aircraft with landing gear before a flight test. In addition, it includes the flight testplanning and the successive phase of analysis required to evaluate the performance of the landinggear and the influence of some design choices.

• Chapter 5: It illustrates the basic features and all the design decisions made towards the prelimi-nary design of the landing gear for FTV16.5 application. It includes all the major outcomes relatedto the conceptual design phases and the initial calculations necessary to define the layout and thestructure of the landing gear, including the design of the required ground test rigs.

• Chapter 6: The basic performance evaluation in terms of heave response of the aircraft with thedesigned landing gear is illustrated in chapter 6. The most critical phases (take-off run and landing),

3


that affects the landing gear design, are considered and compared for both landing gear of FTV7%and FTV16.5%.

• Chapter 7: The last chapter of the document highlights the most important conclusion of thethesis work and the planning of the future works for the design of the landing gear for FTV16.5,expected to be flight tested by the end of 2019.

• Appendix: The appendixes contain the support documentation to understand the features of theMatLab®software for sizing calculations, the CAD assembly drawings and bill of all the componentsused for the 7% landing gear, the theoretical background for the mathematical models used for theperformance evaluation described in chapter 6.

4

Chapter 2

Guidelines for Landing Gear Design

The landing gear can be defined as the essential intermediary that prevents the airplane from the cata-strophe, so particular attention and engineering effort to premeditate possible failure modes, are requiredsince the beginning of the design. [5] Since the functions fulfilled by the aircraft are extremely different,it becomes clear that each landing gear represents an individual case, designed with specific conside-rations and decisions regarding its own application. Several types of landing gear for different aircraftapplications are shown in figure 2.1.

Figure 2.1. Landing gear for different aircraft applications. [6]

In general the following functions are required to the most variety of landing gear system: [7]

• Allow Take-off and landing operations;

• Provide stability for ground maneuvering taxiing and take-off;

• Transfer the ground loads to the airframe;

• Convert the longitudinal kinetic energy in heat thanks to a braking system;

• Damp the vibrations and bouncing caused by the kinetic energy developed upon impact and take-offrun operations;

The undercarriages have essentially to convert the aircraft from its natural airborne environment into alumbering ground vehicle on the ground. The general approach for the design of a landing gear, followsthe normative established by the FAR regulations and typical considerations explained in the pillars ofthe landing gear design like the references Roskam, Currey and Niu. [7] [8] [9] Anyway in most cases thisapproach is not directly applicable for small scaled aircraft, where specific requirements and compromisesbetween scaling process and UAV considerations are necessary.

The three following sections describe the major keys and variables for the landing gear design andsome examples of landing gear designed for airplanes with similarities to FTV 7% and FTV 16.5%, thathave been considered as a reference for some design decisions.

2.1 Design keys and variables for the Landing GearThe design of the landing gear is an iterative process which involves parameters that strongly influencethe aircraft configuration design and aerodynamics performances. [10] All the keys and variables involvedin the landing gear design from the beginning through the whole iterative process, have been evaluatedin the first research phase and are explained schematically as follows.

5


• Scale of the aircraft: It is the first parameter that conditions from the beginning all the designprocess. Essentially the aircraft scale can vary among Radio commanded small planes, UnmannedAerial Vehicles, Ultralight aircrafts, Commercial and Cargo aircrafts, Military aircrafts and for eachof them the specific operation condition of the aircraft highly influences all the design decisions.[11]

• Type and complexity: They represent the two major decisions that must be made before thelanding gear design process can start. The typical possible configurations of a landing gear aretricycle, bicycle, tailwheel or unconventional gear and its complexity is highly affected by the needor not of a retraction system. The optimal landing gear layout can be decided after an analyticalhierarchical process (AHP) method through the assignment of weighting factors to each of thefeature described in the pros and cons table 2.1.

Table 2.1. Pros and Cons of the most used landing gear types [10]

Landing gear typeCharacteristic Tricycle Bicycle Tailwheel

Weight Medium High LowGround stability High Undetermined LowSteering ability High Medium LowLeveled attitude High Medium LowTake-off rotation High Low Medium

The decision to use or not a retractable system is guided by the aircraft cruise speed and weight/costbudget: the state of the arts shows that airplanes with cruise Mach number less then 0.85 tend tohave fixed gears.[7] For light and not fast aircraft applications, the extra weight and the cost thataccompanies a retractable landing gear are usually more disadvantageous than the parasite dragcaused by the friction of the air flowing over the fixed gears. Lightweight airfoil-shaped fairingsand wheel pants can be eventually used to streamline the airflow as aerodynamically as possiblereducing a great amount of parasite drag. [12]

• Landing gear attachment: there are two possible attachment structures for the landing gear,represented by wing and fuselage. Usually the fuselage attachment is preferred for small aircraftswith a fuselage wide enough to allow the desired wheel track. In fact, at equal wheel track, theheight of the landing gear for a wing attachment is bigger (resulting in more weight) and additionalcomponents are needed to transfer the loads from the wing ribs to the fuselage structure. [13]

• Center of gravity position: the center of gravity (C.G.) envelope influences the overall geometryof the system including the horizontal and vertical location of the undercarriages. If the horizon-tal position of the C.G. can vary between wide limits, the worst loading case scenario should beconsidered for the landing gear design. [14]

• Vertical Load ratio: The vertical load ratio between the nose and main landing gear affectsthe position of the undercarriages with respect to the center of gravity as well as the structuralcomponents needed to manage the resulting load distribution. The normal force on the nose gearshould be limited, but not less than 8% of the aircraft landing weight for an adequate steering [7].The usual practice is to design the landing gear in order to distribute the vertical load between8 ÷ 10% and 90 ÷ 92% respectively for the nose and main1 gear. The vertical static load on thenose PN and main gear PM can be calculated according to the system of equations 2.1, where lMand lN are respectively the main and nose arm ratio with respect to the c.g. and ns is the numberof main gear wheel assemblies. [15] {

PN = WlM(lM +lN )

PM = WlNns(lN +lM )

(2.1)

1The "main" gear is so called because it carries the larger amount of load.

6

2 – Guidelines for Landing Gear Design

The nose and main landing gear are even subjected to dynamic loading condition respectively duringlanding with brakes applied and take-off rotation due to acceleration forces developed. The totalnose gear load is therefore obtained adding the dynamic load to the static one, as shown in equation2.2, where ax is the deceleration with brakes applied that is typically 35% of the gravity for a dryconcrete runway. [10]

PNT OT= PN + ax

g

hC.G.W

(lN + lM ) (2.2)

In a similar way the total load on the main gear is obtained adding to the static load, the dynamicload caused by the longitudinal acceleration during rotation, as shown in equation 2.3, where aT isthe average acceleration imposed by the thrust.

PMT OT= PM + aT

g

hC.G.W

(lN + lM ) (2.3)

• Ground clearance: the height of the landing gear should ensures a reasonable clearance betweenthe runway and all other parts of the aircraft in compressed position. The ground clearance requi-rement at take-off, that ensure the prevention of a fuselage or tail hit, is respected if the maximumtake-off rotation angle αTO is less than the clearance angle αc, defined as equation 2.4.

αc = tan−1(Hf

Lf

)(2.4)

The dimensions Hf and Lf , shown in figure 2.2, are respectively the fuselage clearance in leveledposition and the distance aft of the main gear to the beginning of the unsweep angle of rotation.

Figure 2.2. Ground clearance requirement during take-off rotation. [10]

• Take-off rotation and tip-back prevention: the geometry of the landing gear is also affected bythe necessity to have a regular take-off rotation and an adequate tip-back prevention respectivelyduring take-off and landing. The two parameters (αTO and clearance) involved in this designvariable are shown in figure 2.3.

Figure 2.3. Take-off rotation and tip-back parameters. [10]

A regular take-off rotation is ensured if the A angle between the center of gravity position c.g. andvertical line on the ground contact is at least equal to the tip clearance angle αC and higher than15◦. The A angle should not be too much different from the tip clearance angle, otherwise a greatamount of load is needed on the tail in order to rotate the aircraft at take-off. [8]

• Overturn prevention: The overturn of the aircraft on ground is the rolling over of the aircraftthat can happen during ground turning and cross-wind conditions. The phenomena is prevented ifthe moment generated by the aircraft weight about one of the main gear contact point, is higherthan the moment generated by the centrifugal force in ground turning maneuvers and the momentgenerated by acting force in cross-wind conditions. [7] The respect of the requirement passes throughan analysis of ground turning controllability and stability during cross-wind conditions. [10]

7


• Structural integrity: The landing and taxiing loads over rough runways should be absorbed anddissipated through proper struts, possibly with either stiffness and damping properties, capableto maintain as minimal as possible the structural deflection of the landing gear. The maximumdeflection of the landing gear, during impact loading, represents a limit for the maximum track ofthe main landing gear.

• Safety: It is important to consider from the beginning of the design that a failure of any landinggear parts does not represent a risk to critical damage the airframe and system components of theaircraft.[16]

• Low cost and low weight: The use of components off-the-shelf (COTS) is highly encouraged inorder to avoid the cost of custom design whenever possible. If necessary, the landing gear designershould evaluate the possibility to use low cost and low weight components, designed for applicationsthat goes beyond the aeronautical ones.[17]

2.2 Regulations for Landing Gear DesignThe Unmanned Aerial Vehicle applications for the Landing Gear do not always allow the applicability ofstandard regulations. Depending on the scale, the UAV landing gear design can be considered in betweenthe design of the system for Radio command small planes and ultralight aircrafts. The nature of theseaircrafts implies, in some cases, a custom design depending on the specific mission and runway, where theplane is going to operate. Anyway, since the scaled FTVs considered are a representation of an aircraftthat will necessitate a Federal Administration Regulation (FAR) certification, some design choices andparameter definition can be done referring to FAR 23 regulation, valid for normal, utility and aerobaticaircraft with maximum take-off weight less than 12500lbs. [12] The next subsections highlight how toselect some important design parameter and the needed procedures for testing.

2.2.1 Design limit parametersThe most important parameters, that the landing gear designer has to select in a phase of definition of allthe requirements, regard the design landing to which the landing gear structure is expected to respondin elastic field. They are related to the vertical and longitudinal behaviour of the structure, as describedin the following itemize:

• Rate of descent: The rate of descent2 of the aircraft for the landing gear design should be inbetween 7 and 10 fps and can be determined using the equation 2.5, derived from FAR 23.725normative. The values W [lbs] and S [ft2] refers to the aircraft weight and wing reference surfacein landing condition. [18]

wTD = 4.4(W

S

) 14

L

(2.5)

• Ground reaction factor: The vertical dynamic loads developed at the ground contact point onlanding are treated in the first phase of the design as quasi-static loads and obtained by multiplyingthe static load times the ground reaction factor NG, as shown in equation 2.6. A typical value forNG for ultralight aircraft is 3, simulating an impact acceleration of 3G3. [7]

NG = Dynamic loadStatic load (2.6)

• Spin-up/ spring-back loads: In absence of specific tests for determining spin-up and spring-backloads, developed on landing when the wheels pass instantaneously from null speed to the aircrafthorizontal landing speed, the appendix D of FAR 23 regulation can be used for an initial estimation.

2The rate of descent is often called sink speed or touchdown rate and is defined as the vertical speed of the aircraftbefore touching down. It is dependent on the flaring speed and angle of the aircraft in a landing attitude of two pointscontact on the two main landing gears.

3The impact acceleration is often measured as function of the gravity acceleration, so 3G corresponds to 29.43m/s2

8

2 – Guidelines for Landing Gear Design

The maximum value for the horizontal force (in pounds) acting on the wheel is determined throughthe equation 2.7,

FHmax = 1re

√2Iw (VH − Vc)nFVmax

ts(2.7)

where re is the effective rolling radius (in ft) of the wheel under impact, Iw is the rotational massmoment of inertia of rolling assembly (in slug ft), VH and VC are respectively the linear horizontallanding speed of the airplane and the peripheral speed of tire (in fps) if pre-rotation is used, n isthe effective friction coefficient4, FVmax

is the maximum vertical force on the wheel (in lbs), ts5 isthe time interval between ground contact and reaching of the maximum vertical force on the wheel(in seconds).The dynamic spring-back effect can be estimated, in a level landing condition, assuming the loadin equation 2.7 to be reversed.

2.2.2 Testing proceduresThe FAR regulation contains also the definition of some testing procedures needed to certify the designof the landing gear. The most relevant test peculiarities are presented as follows:

• Aircraft attitude: For a leveled landing the attitude of a tricycle aircraft should be one of thefollowing:

– Simultaneous nose and man wheels contact on the ground;– Nose wheel clear of the ground when main gear touches down;

• Limit Drop Tests: The full airplane or equivalent assemblies consisting of wheel, tire and shockabsorber should be drop tested from free drop height in inches not less than 9.2 inches and notmore than 18.47 according to the equation 2.8, derived from FAR 23.725 regulation. [18]

hDT = 3.6(W

S

) 12

(2.8)

The drop weight to use in equivalent drop tests should be determined by the equation 2.9.

We = W[hDT + (1 − L) d]

(hDT + d) (2.9)

where hDT is the drop height calculated with equation 2.8, d is the deflection under impact of thetire plus the vertical component of the axle travel relative to the drop mass, L is the Lift to weightratio6, W is the aircraft landing weight or the static load on the main gear WM or the static loadon the nose gear WN if the drop tests are done considering assembly units. The limit inertia loadfactor n, applied to the center of gravity c.g., should be determined by the equation 2.10.

n = njWe

W+ L

W(2.10)

where We and W are respectively the equivalent weight of drop test and the aircraft landing weight(or the static weight on the main gear or nose gear, if the drop test is done with equivalent assemblyunits), and nj is the load factor recorder in the drop test ((dv/dt)/g) plus 1.

• Shock absorption Tests: the limit inertia load factor in 2.10 selected for the design should notbe exceeded in energy absorption tests. The test must demonstrate the landing gear not to fail ina simulated descent velocity equal to 1.2 the selected rate of descent in equation 2.5, assuming thewing lift equal to the aircraft weight before the impact.

• Tire rating Tests: Each tire should have a tire rating not exceeded by corresponding static groundreaction under the design maximum weight and critical position of the center of gravity.

4A typical value of the friction coefficient on landing is 0.80. [18]5A typical value of the time delay from the contact point on landing and the attainment of the maximum vertical

load is 0.2s. [19]6The Lift to weight ratio should be less than 0.667 according to FAR 23.725.[18]

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2.3 State of the art of Landing Gear for similar aircraft appli-cations

The present section illustrates the state of the art of landing gear for different aircraft applications, fromsmall radio commanded planes and Unmanned Aerial Vehicles to ultralight aircraft certified with FARregulations. Some of these aircrafts, with weight and performance similarities to FTV 7 and FTV 16.5,have been considered as a reference to select the off-the-shelf components and to make design choices.Figure 2.4 shows the weight and speed range of the aircraft applications considered, and where the twoblended wing body FTVs are located among four of representative aircrafts for each category.

Figure 2.4. Landing gear application for several UAVs

2.3.1 References for FTV7%The 7% scale of the new generation blended wing body aircraft is located in between RC commandedhobby airplanes and small jet UAVs. The Skymaster RC jets represent the framework of tires and wheelsoff-the-shelf solutions that can be considered for aircraft with gross weight and speeds similar to FTV7.The tires for this aircraft application, are designed to withstand 40 m/s landing and take-off speed. [20]

The tricycle landing gear of the Penguin C has been considered as a reference for designing thecustom leaf strut and the nose gear strut for FTV7. The suspension of the landing gear is essentiallyattributed to the elastic behaviour of the bended aluminum main strut and to the simple spring elementsinside the nose gear strut. The steering system is implemented using a servo linkage mechanism withself-centered caster nose gear strut. [21] The aircraft, provided at the Center for Aerospace Research, hasbeen considered even as a reference for the ground mock-ups necessary to tune the control parametersfor a fully operative autopilot on an aircraft with undercarriages.

2.3.2 References for FTV16.5%The larger flight test scale demonstrator can be considered part of the experimental UAV category, wherethe use of specific custom components, designed for the required mission, becomes highly discourageddue to the high cost and time required. For big aircraft scales, it is necessary to add some form ofshock absorber to better manage dynamic load conditions during take-off and landing, and provide theaircraft with adequate steering and braking system. The general guideline for this category of aircraft isto search for low cost components off the shelf, as implemented on the landing gear of the X-48 BoeingNasa, characterized by shock absorbers and brake system from bike applications. [17]

As for tire and wheel solutions, the choice among hobby aircraft applications gets rapidly limitedabove 150mm diameter. Therefore a reference is represented by nose or tailwheel applications for smallultralight aircraft, such as the Thatcher CX4, even if in most of the cases they are not rated for highspeed due to the low speed requirement for their application.

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Chapter 3

Landing Gear Design andDevelopment for FTV7%

The present chapter shows the detailed experimental design and development of the landing gear systemfor the 7% scale of the blended wing body new generation aircraft. Since the beginning it has beenfundamental to consider the behaviour of the aircraft, already flight tested without undercarriages at theCenter for Aerospace Research, during the critical phases that influence the landing gear performances.An initial description of requirement list, useful to define the framework in which the landing gear designershould operate, is followed by the explanation of all the fundamental design steps for both main and noseundercarriages. All the peculiarities and challenges, linked to the production phase and integration ofthe full system with the aircraft, are presented as well as the ground testing, necessary to validate thedesign itself.

3.1 Requirement list

The first step required to start the design of a landing gear for an already built aircraft is to define a list ofcore requirements, including the constraints imposed by the systems that cannot be significantly modified.The landing gear requirements for FTV7 have been determined considering all the general guidelinesdescribed in chapter 2 and considerations resulted from continuous discussions with the stakeholders.

Five categories of essential requirements have been selected to guide the iterative design process: gene-ral requirements for the landing gear, aircraft performance requirements, scaled-geometry requirementsand specific requirements for the main gear and nose gear. Due to the interdisciplinary and iterativenature of the design for aircraft systems, most of the requirements are interdependent and for this re-ason they have been organized in an Excel sheet, then imported in Matlab, using a parametric designapproach. [22]

In this phase, it has been useful to consider the real behaviour of the aircraft, already flight testedwithout undercarriages several times at the Center for Aerospace research, during landing approach andtouch-down1, in order to evaluate in which ranges of impact the aircraft is used to operate. For thispurpose, the StatData feature of the software Flight test data processing, developed by Sean Bazzocchiand Jenner Richards at the Center for Aerospace research, shown in figure 3.1, has been helpful tocompute the average and maximum values of the parameters related to the impact, that are the rate ofdescent, landing speed and vertical deceleration.

1The phase of take-off did not influence the design of the landing gear since the aircraft was catapulted into theair, and so the resulting accelerations and oscillations were related to the shooting of the catapult.

11


Figure 3.1. Interface of the StatData feature of the Flight Test Data Processing software

3.1.1 General requirements

The most important general requirements of the landing gear design for FTV7 are summarized in table3.1.

Table 3.1. General requirements for the landing gear design for FTV7

# General Requirements Value/Type1 The type of landing gear should be as cheap and simple as possible Tricycle, Fixed2 The aircraft should be able to perform in concrete not well prepared

runwaysMerrit airstrip

3 The aircraft should have an adequate system to facilitate ground ope-rations

Steering system

4 The braking system is not required but can be implemented to test thebraking control parameters with autopilot

Electro/magneticbrakes

5 The landing gear should be the same for different aircraft configurations:FTV7-tailed and FTV7-no-tailed

As described in section 2.1, one of the most important factors that determines the type of the landinggear is the configuration of the aircraft. The flight test vehicle considered, representation of a new blendedwing body concept aircraft, is developed in two different configurations, with tail and without tail. Themain goal for the landing gear designer, in this specific case, is to guarantee the same landing gear betweenthe two configurations in order to reduce the cost and complexity of the design.

3.1.2 Aircraft and Performance requirements

The requirement related to the aircraft and performance are shown in table 3.2. The reason of the largeexcursion of the center of gravity (C.G.) for the take-off and landing operations comes from considerationof the full scale aircraft: the blended wing body, with and without tail, is supposed to fly in non stableregimes and this implies the necessity of shifting the position of the C.G. to trim and balance the aircraftin all the flight phases.

The data values of performance used for design considerations are the one obtained from the previousflight tests of FTV7 without landing gear, using the statistical software described in section 3.1. The

12

3 – Landing Gear Design and Development for FTV7%

maximum values, especially as for the accelerations registered by the 50 Hz log file2, are referred to themaximum of the peak values for each flight, that in most of the cases are developed in a short timespanlike 0.1 seconds. For this reason, they are useful to have a full picture of all the possible worst casescenarios experienced by the aircraft on landing, but nor directly applicable as design parameters, betterrepresented by average values.

Table 3.2. Aircraft and performance requirement for FTV7

# Aircraft performance requirements Value/Type1 Maximum Take-Off weight of FTV7 with tail 13.6 kg2 Percentage of the center of gravity position for take-off and landing with

respect to the mean aerodynamic chord (M.A.C.)56 ÷ 66%

3 Average Landing Speed from previous flight tests 26.80 m/s4 Maximum Landing Speed from previous flight tests 30 m/s5 Average Rate of descent from previous flight tests 1.05 m/s6 Maximum Rate of descent from previous flight tests 1.4 m/s7 Average landing deceleration from previous flight tests 1.1G8 Maximum landing deceleration from previous flight tests 4.3G

3.1.3 Geometry requirementsThe geometry of the landing gear should replicate as close as possible the limits imposed by the fullscale aircraft. The most relevant geometry requirements are shown in table 3.3. The position of the nosegear is fixed to the scaled value from the full scale aircraft, whilst the location of the main gear can bemoved depending on the center of gravity position. The design choice for the main gear position has beendone in order to have a load distribution of 90% on the main gear and 10%, verifying that each locationrespects the take-off rotation and tip-back criteria introduced in section 2.1.

The vertical behaviour of the landing gear is influenced by the definition of the height of the landinggear in all the phases that involves the landing: fully extended, static, fully compressed position whererespectively no loads, MTOW corresponding loads and impact loads are applied. The static position isdetermined by the ground clearance criteria while the fully compressed position should replicate possiblythe scaled geometry.

Table 3.3. Geometry requirements for FTV7

# Geometry Requirements Value/Type1 Position of the nose landing gear with respect to the front fuselage

station275 mm

2 Load ratio on the landing gear (nose, main) due to the C.G. position 10%-90%3 Main gear track measured at the outboard wheel position, in static

position of the landing gear589 mm

4 Vertical height of the gear in extended position with respect to the waterline (W.L.) reference

174.89 mm

5 Scaled compression of the landing gear from its extended position 27.26 mm

3.1.4 Main and Nose gear requirementsSpecific requirements for main and nose gear are presented in table 3.4. In a tricycle configuration themain and nose gears have the same height, so the aircraft is leveled on ground even if the main geartends to have bigger wheels. The maximum allowed size with respect to the scaled diameter, has beenselected to be 25% larger, after some considerations that consider the wheel’s market among RC planes

2The log file at 50 Hz is installed on the aircraft and records all the flight data used from the autopilot every 0.02seconds.

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for comparable applications to FTV7 and maximum allowed drag penalties due to the increased frontalarea, in comparison with the values expected for the full scale aircraft. [3]

Table 3.4. Specific requirements for the main and nose gear of FTV7

# Requirement Value/Type1 Scaled tire size for the main landing gear 2.5 inches2 Scaled tire size for the nose landing gear 1.5 inches3 Maximum larger percentage for tire sizes 25%4 Number of wheel assemblies per each strut 1

All the stated requirements have been imported in a developedMatlab® parameter calculator software,described in Appendix A, in order to define the resulting feasible geometry and preliminary layout of theaircraft with undercarriages.

3.2 Design and development of the main gear assemblyThe present section describes the process of design and integration of the main landing gear for the7% scale of flight test vehicles, in both tailed and no-tail configurations. The design has moved rapidlyfrom scratch to solid CAD modeling in SolidWorks and the final modeled layout, with all the neededcomponents, is shown in figure 3.2. The choice of the off-the-shelf components and the custom design ofthe leaf strut, as well as the features of the main landing gear, are explained in the following subsections.

Figure 3.2. CAD assembly of the designed main landing gear for FTV7

3.2.1 Main gear features

The main landing gear has been designed in order to support all the maximum take-off weight of FTV7,that simulates the case of not perfect three points landing. The initial approach has been to possiblysearch for off-the-shelf components in order not to affect drastically the cost and weight budget of theBombardier-Quaternion project. The final assembly of the main landing gear, shown in figure 3.2 consistsof wheels derived from the Skymaster radio commanded jets, introduced in section 2.3 and customcomponents for the leaf strut and the parts necessary for the integration with the airframe.

The most interesting feature of the main landing gear is the possibility to move easily its locationdepending on the center of gravity position. This capability has been required in order to test the workingoperation of the autopilot during take-off and landing, varying the longitudinal C.G. envelope from 56%to 66% of the mean aerodynamic chord (M.A.C.). [2] The main landing gear is attached to the flat partof the fuselage, using apposite shifting slots, and it is in part covered by the redesigned internal foam ofthe belly pan.

The structural stiffness of the main landing gear is carried by the elastic behaviour of the strutleaf, whilst the damping of the system is developed in the most part by the pure inertial motion of theaircraft and is slightly increased by using additional rubber cushion sheets in between the landing gearattachments and the skin of the airframe.

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3.2.2 Wheels and Tires for the main gearThe wheel assembly is the component responsible to allow the aircraft to have contact with the ground.For small aircraft applications it is usually a nylon or metal alloy wheel, that houses the bead seat for afoam or rubber tire. Several options have been considered in a preliminary market research, all suitablefor the size and weight of FTV7. The most interesting solutions, shown in figures 3.3 and 3.4, have beenpurchased and characterized.

Figure 3.3. Option 1: Sky-master wheels

Figure 3.4. Option 2: TurnigyHK wheels

The final decision has been done based on an analytical hierarchical approach with weighted features de-cided in relation to the potential application to FTV7, that is summarized in table 3.5. Since the aircraftis supposed to operate in runways with a considerable length, it is expected to be stopped without theneed of a braking system. In addition, no fairing have been designed and so the parasite drag caused bythe rolling frontal area of the wheel is expected to provide a backward force that can help to stop theaircraft.

The option 1 has been selected because it privileges the weight saving and simplicity. However if theairplane operation will move to short and unprepared airstrip, a braking system should be essential, andso the electro-brakes of the option 2 would be preferable. The electro-magnetic brakes are easy to controland to integrate with the overall system: when the current is passed through the coil, the magnets intothe outer rim of the wheel react with the magnetic field generated and slow the wheel down, thus causingthe aircraft to brake. As there are no friction parts in the wheels, they cannot wear out so reliability andlongevity should never be an issue. This solution could potentially be used even to brake progressively,avoiding the wheels to lock-up, and differentially between the left and right wheel, helping the aircraftto steer. [23]

Table 3.5. Analytical Hierarchical table for tire and wheel selection of FTV7 main landing gear.

Option 1 Option 2Features Value Rank Value Rank

Weight (50%) 83 gr 10 192 gr 5Size (20%) 2.7"x0.8" 8 3"x0.8" 6Brakes (5%) Extra 2 Electro-Magnetic 8

Application (25%) Model: 15 ÷ 20 kg 8 Model: 10 ÷ 15 kg 6Total 100% 8.7 5.6

3.2.3 Design of the main gear strutThe design and development of the main gear strut started from the evaluation of possible off the shelfcomponents and moved towards a custom solution, including CAD design, production and testing. Thissubsection describes in detail the design and production processes performed, as well as the simulationand ground tests done to characterize the component and verify the design decisions made.

Off-the-shelf solutions

The impact loads for such a small scale of airplane can easily be manageable by a strut leaf in Aluminum orCarbon fiber composite. Several off the shelf components, designed for radio commanded small aircraft, instock at the Center for Aerospace Research, have been tested and evaluated. For the sake of brevity onlythe structural tests of the Aluminum 2024 leaf with thickness 1/8” are presented, since it has the mostsuitable geometry in comparison with the requirements stated in subsection 3.1.3. The landing scenario

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simulated by structural simulations and ground drop test, shown in figures 3.5 and 3.6, represents animpact of 3G and all the equivalent MTOW weight acting on the strut. The landing condition hasbeen simulated applying the procedures described by FAR regulation, introduced in subsection 2.2.2, andconsiderations from the previous flight tests without undercarriages.

As result from both simulations and ground tests, the component has been deformed in plastic field,therefore it has not resulted stiff enough to be used for the FTV7 case design.

Figure 3.5. SW structural simulations of theAl2024-1/8" leaf strut

Figure 3.6. Drop Test of the Al2024 1/8"performed at CfAR

Due to the difficulty to find a leaf strut among radio commanded applications, with the exact geometryrequired for the fully extended and fully compressed positions, a custom solution has been planned to bedesigned and developed. A solid model with the same geometry as the off-the-shelf component, has beenevaluated in SolidWorks with different Aluminum grades and thickness, in order to guide the choice ofmaterial for the custom design. The simulations suggested to improve the grade of Aluminum to ERGAL7075 and the thickness of almost 1.5 mm, without increase considerably the weight of the component.

CAD design of the custom leaf strut

The design phase of the leaf strut has been a trade-off among the Coward methodology described inreference [8], the requirements stated in section 3.1 and the necessity to integrate the strut with theairframe of FTV7 without internal and external interference. The most important design parameters areillustrated in figure 3.7, where L is the equivalent geometric arm between the force on the wheel and thetop strut bending radius, t is the thickness of the strut, αC , αs and θ are respectively the camber, sweepand bending angles of the strut, WR and WB are the evaluation of the strut width at the two bendingradius locations.

Figure 3.7. Geometric parameters for the leaf strut design

The values of WR and WB can be calculated from the graph used in the Coward procedure [8], shownin figure 3.8. The root width dimension can be estimated entering in the graph with the beam widthparameter b, calculated according to the experimental equation 3.1, valid for Aluminum alloy struts,

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where W is the aircraft weight and Sref is the reference wing area.3 The width at the bottom bendingradius of the strut is assumed to be half of WR, and the thickness should be about t = WR/8.

Figure 3.8. Design taper para-meters for the leaf strut. [8]

Figure 3.9. Integration of the main gear with themid bay of the aircraft

b = 0.01373W 1.5

LSref0.5 (3.1)

The θ angle is defined by the bending of the strut, necessary to guarantee the required vertical positionof the landing gear. The wheel travel of the landing gear is then influenced by the camber angle αC ,determined in an iterative process in order to have the wheel front section aligned with the vertical linewhen the strut is loaded in a 3G impact landing with maximum take-off weight. The sweep angles at thefront an rear location of the strut leaf leg have been dictated by the integration of the component withthe airframe, since the connection is supposed to be inside the mid bay of the aircraft, as shown in figure3.9. In addition, a swept leaf strut is expected to have a better stress distribution during impact, sincein most of the two points landing on the main gear with the nose gear clear from the runway, the groundreaction results to be transferred vertically to the strut leaf.

The resulting design geometric strut parameters are organized in table 3.6.

Table 3.6. Geometric parameter determination of the custom leaf strut design

t θ αC αs−f αs−r L b WR WB3/16" 31.33◦ 7◦ 9.4◦ 2.1◦ 6.55" 0.075" 3" 1.5"

Material procurement and production of the customized strut

The first step of the production process has been the procurement of the material sheet in Aluminum7075-T6, with thickness 3/16" and size 24"x24". The size has been selected in order to produce two setsof main gear strut for each aircraft FTV 7tailed and FTV 7no−tail and one extra spare part for testing.

The flat pattern of the component has been obtained from the material sheet using abrasive waterjetcutting technique, selected because it is a cold process that makes no impact on the material being ma-chined. At its most basics, a slurry of water, abrasive and air, flows from a pump with pressure between60000 and 94000 psi, through plumbing with a cutting head of minimum kerf diameter around 0.035inches. At the cutting head, a high speed air valve, allows the water to pass through the jewel orificecreating a supersonic waterjet stream, typically at Mach 3.0, able to cut the material without create heataffected zones (AHZ) typically caused by other machine processes that require for instance warping andclamping.[24]

The next step of production has been to obtain the final geometry through the bending process ofthe water-cut parts. Appropriate bending brakes have been used to lock the part and load force around

3All the data in the beam width parameter equation are espressed in British imperial units.

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the bending lines. The initial attempt to bend the part has been done as usual for the bending of lowerAluminum grades, with a cold process machine. As a result the component has started to crack at thelocation of the bending line, as shown in figure 3.10, due to the hardness of the 7075 Aluminum, the largethickness and sharp bending radius 1/2". For this reason, it has become necessary to anneal the bendinglines by heating at temperature4 around 400 − 450◦C in order to bend properly the parts.[25] The resultof the hot bended leaf struts is shown in figure 3.11.

Figure 3.10. Crack on the lower bending ra-dius of the leaf strut, after the first bendingmachine process

Figure 3.11. Final bending hot process ofthe main gear struts

After machine processes it is important that the material recover ideally the original properties and so atypical heat treatment for the Aluminum 7075 has been applied, following the specifications ASM 2770-2658 [26]. The leaf strut has been first heat treated at temperature 870◦F for 1 hour and 11 minutes,then it has been quenched for 7 seconds and finally the precipitation age for curing the material has beenprocessed at the temperature of 250◦F for 23 hours and 10 minutes. The process, in terms of temperaturewith respect to the time intervals, is shown in figure 3.12.

Figure 3.12. Heat treatment of the leaf strut. [27]

3.2.4 Testing of the main gear

The Main Gear leaf strut has been tested firstly with finite element (FE) simulations in SolidWorks andthen ground tested by means of Static tests to estimate the structural properties and Drop tests to verifythe integrity of the component when all the aircraft weight is acting on the main gear in case of 3Gimpact landing.

4Aluminum 7075 should be formed at temperature not too close to the melting point that is 477 − 635◦C, since itwould start to get brittle.

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Static tests for the main gear strut

The theoretical linear stiffness can be calculated using the equation 3.2, described in the Coward metho-dology in [8], considering the Young Modulus of Aluminum 7075 equal to 10400 ksi.

Kth = E ·W 3R

96b · cos θ2 (Lt

)3 = 32.067 N

mm(3.2)

The corresponding analytical displacement can by estimated considering one of the leg of the strut leafto be approximated by a rectangular beam, using the equation 3.3.

zana = FM · cos (θ) · Leq3

3EIeq= 4FM · cos (θ) · Leq3

E · beq · t3(3.3)

The beam approximation implies to consider equivalent geometric parameters to the ones of the realtapered strut described in figure 3.7. The equivalent beam parameter beq is defined as the average of thestrut width at the root WR and bottom WB locations, whilst the equivalent length of the beam Leq hasbeen derived assuming to have the same surface area between the real strut and the approximation. Theother parameters are the same described in table 3.6.

The same structural properties have been estimated using structural FE simulations in SolidWorks,with increasing weight, from 0 kg to 30 kg with 3 kg of each step. The boundary conditions used for thefinite element simulations, that try to represent the reality as close as possible, are bearing constraints atthe attachment points with the airframe and remote loads at the axle location of each leg. The geometryis modeled with 10557 solid tria mesh elements with overall size length 4.56 mm, capable to predict thebending behaviour of the material.

Figure 3.13. Displacement of the main gar strut with a static load corresponding to the MTOW of F T V 7

For sake of brevity, only the loading condition corresponding to the static load distribution of themaximum take-off weight of FTV7 is presented as follows. The corresponding maximum computationaldisplacement, shown in figure 3.13, is 3.4 mm registered at the same location of application of the remoteloads.

The tests, with the same increasing weight, have been replicated in reality at the shop of the Centerfor Aerospace Research, by using a simple drop leverage test rig and digital caliper measurement for thedisplacement. The measurements have been taken between the ground reference and a reference on theTest rig, in both loading and unloading case. The material behaviour has been proved to be elastic, since

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the outboard track between the axles at the end of the tests has remained the same of the unloaded case.The experimental maximum displacement, under equivalent maximum take off weight load, is proved tobe similar to the computational one (3 mm) , as shown in figures 3.14 and 3.15.

Figure 3.14. Test rig with no loading appliedFigure 3.15. Test rig with static loading cor-responding to the MTOW of the aircraft

The experimental (zexp), computational (zcomp) and analytical (zana) displacements resulting fromthe increasing static weight up to 30 kg are summarized in table 3.7.

Table 3.7. Resulting displacements from Static Tests with increasing weight: (0 : 3 : 30kg)

# Test Load [kg] zexp[mm] zcomp[mm] zana[mm]Loading Unloading

1 3 0.31 0.2 0.76 0.922 6 0.97 1.38 1.53 1.843 9 1.76 2.01 2.29 2.764 12 2.55 2.69 3.06 3.685 15 3.09 3.50 3.82 4.596 18 4.45 4.57 4.58 5.527 21 5.64 5.75 5.35 6.448 24 6.20 6.40 6.12 7.369 27 7 7.2 6.88 8.2810 30 7.75 7.90 7.64 9.20

The difference in some values between loading and unloading condition can be justified due to thedifficulty to take the measurements each time between the same points and to the not exactly sandequivalent weight used for each test. Due to the measurement uncertainty an average between the twoloading and unloading cases can represent a better estimation of the real behaviour.

The results of table 3.7 are shown graphically in figure 3.16, where the displacements are approximatedwith polynomial trend for the experimental case and linear one for the analytical and computational cases.In the reality the strut stiffness, represented by the slope of the curve, progressively decreases with theincreasing of the load demonstrating that the real behaviour is not linear.

A double least squares approximation can be used to linearize the real stiffness properties of the strut,in two different ranges of loads, denominated 1st slope (with load less than 170 N) and 2nd slope (withload more than 170 N). In the first load range the real linearized stiffness results to be well estimated bythe computational analysis and much more stiff respect to the analytical model: this is probably due tothe assumption used to approximate the tapered strut leaf as a single piece of rectangular beam.

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Figure 3.16. Load displacement curve for the main gear leaf strut

In the second range of loads, the strut progressively loses stiffness due the geometric non linearities ofthe component: in this case the structural properties are better estimated by the analytical approach withrespect to the computational one, were inconsistencies are registered, probably due to the not uniformdistribution of vertical load on the wheel axle when the load increases. The resulting values of linearizedstiffness in all the cases analyzed, with the respective error eR respect to the real linearized experimentalbehaviour, are described in table 3.8.

Table 3.8. Linearized stiffness from experimental, analytical and computational tests

Exp 1st slope Exp 2nd slope Analytical ComputationalK [N/mm] 40 30 32.067 38.58

eR 1st slope X X −20% −4%eR 2nd slope X X +6.5% +22.2%

Impact drop test of the main gear strut

The Drop tests for the main gear have been performed according to FAR 23 regulations, described insubsection 2.2.2, and considering a total weight corresponding to the MTOW of the aircraft in tailedconfiguration, the worst case scenario of a two points landing on the main gear. The tests have beenperformed in the shop of the Center for Aerospace Research using the same test rig of the static tests,shown in figure 3.14.

Different landing cases have been simulated, as shown in table 3.9, considering as variable input thevertical rate of descent and the drop height calculated according to equation 2.8. The first three casesreproduce typical landings experienced by the previous flight tests of FTV7 without undercarriages, withan average impact acceleration that is in between 1.8 and 2.4 the gravity. The last landing case, is thedesign landing corresponding to an average impact acceleration of 3G: the resulting displacement (27mm) is quite similar to the scaled down compression of the main landing gear, stated in the requirementof subsection 3.1.3, demonstrating that the main landing gear has been designed in structural similitudewith the full scale aircraft.

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Table 3.9. Drop test setting and results for different simulated landing cases

SinkSpeed

Dropheight

Compression Time tostop

Impact acce-leration

LandingCase 1

1 m/s 51 mm 14 mm 28 ms 1.8G

LandingCase 2

1.16 m/s 69 mm 16.6 mm 28.6 ms 2G

LandingCase 3

1.5 m/s 114 mm 24 mm 32 ms 2.4G

LandingCase 4

1.81 m/s 167 mm 27 mm 29.8 ms 3G

3.3 Design and development of the nose gear assembly

The second, but not less important, component of a tricycle landing gear is the nose (often referred asfront) undercarriage assembly. The design, procurement and production of all the nose gear components,has followed the same general approach of the main gear, evaluating and testing off the shelf componentswhenever possible and developing custom parts to enable the operation and integration with the aircraft.The CAD of the front gear system is shown in figure 3.17. All the relevant components shown, arediscussed in detail in the following subsections.

Figure 3.17. CAD assembly of the designed nose gear for FTV7

3.3.1 Nose gear features

The nose gear functions for FTV7 are basically accomplished by an assembly of aluminum parts, helicalsprings inside the nose strut and rubber tire, providing the required support, stiffness and dampingproperties. The steering ground maneuverability of the aircraft is carry out by an electric servo thattransfer the electric power in motion to the front wheel thanks to a shaft-coupler, horns and mechanicallinkage. Custom made components have been designed and manufactured to attach the nose landinggear to the airframe and provide the needed kinematics and mechanics for the steering maneuvers.

The fixed location of the front wheel is defined by the document for requirement specifications inreference [2] and, as a consequence, the nose landing gear results to be loaded vertically by 10% of thethe total aircraft weight.

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3.3.2 Component off-the-shelf for the nose gear

The off-the-shelf components used for the nose landing gear are essentially the nose strut and wheelassembly, both selected after evaluation and testing of several solutions from RC jet applications. Dueto the impossibility to fully integrate the components, designed for other specific small planes, in mostof the cases it has been necessary to edit the parts in order to increase the reliability and the integrationwith the airframe.

Strut

The selected nose gear strut, in Aluminum 6061, has been designed as main gear leg for scale models asthe Warbird jets. It consists of three different parts: piston, cylinder to house the compression springsand a double scissor to connect the two structural parts and allow the motion in loading condition. Todemonstrate the possibility of using that solution for the nose landing gear of FTV7, it has been droptested according to the regulations already used for the main gear testing, stated in section 2.2.2, with anequivalent weight corresponding to 10% of the aircraft maximum take-off weight. The stiffness propertiesare attribute to the internal springs, adjusted in order to have a double compression behaviour dependingon soft or hard landings. The static position, with static load applied on the strut, and bottoming out ofthe strut are shown respectively in figures 3.18 and 3.19.

Figure 3.18. Nose strut with sta-tic load corresponding to the aircraftMTOW

Figure 3.19. Nose strut with botto-ming force, corresponding to the im-pact in hard landings

Due to the necessity to respect the height constraints stated in subsection 3.1.3 and to integrate the com-ponent with the internal front bay’s space of the aircraft, it has been necessary to edit the cylinder of thestrut reducing the length of the leg by 10 mm. An additional aluminum bushing has been manufacturedand insert to reinforce the strut where there is the accommodation for the wheel axle.

Wheel assembly

Aluminum wheels and rubber tires, designed for mid-sized turbine RC jets, have been selected to beapplied for the nose landing gear. The wheel axles have been provided with stainless steel flanged ballbearings with rating ABEC-4 and capable to withstand easily the load acting on the tire.

3.3.3 Design of additional custom components

Additional components for the nose gear integration and for the implementation of the steering systemhave been designed and manufactured, using the mechanical machines provided in the shop of the Centerfor Aerospace Research and in the mechanical laboratory of the University of Victoria. The followingparagraphs show the CAD design and production process of the needed custom components, illustratedin figure 3.20.

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Figure 3.20. CAD design of the custom components for the nose gear of FTV7

Attachment plates

The Aluminum 7075 attachment plate has been designed in order to easily integrate the nose gear withthe airframe and provide the needed alignment for the shaft of the steering system. The design goal hasbeen to install the nose landing gear at the same location where the launcher hook for the catapult systemwas bolted, in order to reduce the amount of extra connections to the aircraft. The component, as shownin figure 3.20, has four counter-sunk clearance hole for the connection 6-32 screws with the airframe, twotap screw holes for the spring screws 2-56, and four tap screw holes 6-32 for the attachment of the strutcylinder. Two additional buttonholes have been placed in the central area of the component in order torelease material where it is not for structural support. The component has been manufactured using thesame material of the main gear leaf strut and the same waterjet technique, already described in subsection3.2.3.

An extra carbon fiber plate with thickness 1/16", has been used in between the skin of the aircraftand the Aluminum plate, with the scope to mate the attachment plate to the aircraft as uniformly aspossible, in the same location where the launcher hook was installed and so where the skin surface wasnot perfectly smooth.

Horns for the steering system

The steering motion is transferred from an electric servo, positioned inside the front bay, to the nose strut,through a servo horn and a strut horn, shown in figure 3.20, respectively clamped to the servo shaft andto the nose gear piston strut. The design of the two parts has been iterative, in order to facilitate themounting procedure in a limited space as the front bay of the aircraft and to reduce the amount of torsionload, needed on the screws to clamp the horn to the shaft and strut cylinder: the clamping procedureis simplified thanks to the design of rectangular houses with edges parallel to the faces of the hex nut,resulting in no rolling tendency of the nut, when subject to torsion.

The final design of the strut horn has a slot in the internal diameter, in order to allow the screwthat holds the nose strut piston to pass through when the load on the wheel increases. Another featureof this part is the possibility to be connected, with extension springs, directly to the aluminum plate,maintaining the wheel self-centered in the mid-line axis even in case of servo or linkage failures.

Cylinder for the nose gear strut integration

The nose gear strut is connected to the airframe by using an Aluminum 2024 cylinder that passes throughthe skin of the aircraft and is bolted to the attachment plate, as shown on the right of the figure 3.20. Thetranslational and rotational motion of the strut, along longitudinal and transversal direction, is preventedthanks to the use of delrin sleeves around the strut, while the vertical location is hold by a lock-tidedscrew on the top of the cylinder, rotating on a delrin washer when the strut is steered.

3.3.4 Design of the steering systemThe steering system has been designed to transfer the motion from an electric servo to the front wheelassembly thanks to a mechanical linkage. The steering ability of an aircraft is measured in terms of the

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maximum steering angle and consequent turning radius, necessary to perform a 180◦ turn within thewidth limit of the runway. [28] This requirement is less restrictive for small UAVs, where usually therunway of application is much more wide than the actual wing span of the aircraft.

The maximum steering angle of the nose gear for FTV7 has been assumed to be ψ = ±25◦, the samelimits of the rudder control already tuned in the autopilot. The basic geometry and kinematics of thesteering system is shown in figure 3.21. The turning capability of the aircraft is evaluated based on theturning radius Rturn corresponding to the maximum steering angle ψ, as described in equation 3.4 andshown in figure 3.22.

Figure 3.21. Geometry and kinematics of thesteering system

Figure 3.22. Turning radius for thesteering system

The parameter lM−N is the wheelbase and ψ is the maximum steering angle, measured with respectto the longitudinal centerline of the airplane.

Rturn = lM−N

cos(90◦ − ψ) (3.4)

The three bar mechanism of the steering system has been designed in such a way to ensure themaximum steering angle when the servo arm has an inclination of β = ±30◦, in order to not haveinterferences with other internal components. The resulting arm ratio is defined as in the equation 3.5.

hstruthservo

= tan (ψ)tan (β) (3.5)

All the components used in the steering system, including standard parts (electronic servo and mechanicalbar) and custom parts are shown in figures 3.23 and 3.24. Due to internal space constraints, only theservo is allocated inside the nose bay of the airframe, whilst the other components of the steering systemare positioned under the flat belly of the aircraft.

The servo HV6130-MKS has been selected among RC airplane series applications and it results to bethe best compromise among size, weight and torque required to steer the wheel. It has an aluminum caseand metal gear, capable to transfer electric power in mechanical motion with very low slug and adjustablespeed. [29] The attachment support for the servo has been designed to be produced with 3D printingtechnique in versatile plastic. The part has an inclined surface with branched ribs in order to reinforcethe support and facilitate the bonding with the front wall of the nose bay of the airframe. On the otherside, it allows the mating with the servo replicating its detailed shape. An adequate clearance hole allowsthe servo gear to rotate and move the shaft without interferences with the support.

The motion is transferred to the other components under the flat belly thanks to a shaft connectedto the servo with a coupler-screw mechanism and passing inside a delrin bushing through the airframe.The mechanical transmission is granted by connecting the linkage to both servo shaft horn and nose struthorn.

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Figure 3.23. Steering servocomponents Figure 3.24. Mechanical steering system components

Two additional extension springs are used as backup system in case of failure of the servo mechanism,that can occur during hard landing or high irregularities of the runway. The springs connect the struthorn to the attachment plate and they have been selected in order to allow the maximum design steeringangle ψ. The two springs work oppositely in extension and compression, maintaining the front wheel inthe neutral position when no motion is applied form the steering system.

The spring stiffness has been selected considering an equilibrium state about the center of the wheel,illustrated in figure 3.21, when the limit positions for the steering system are reached with steering anglesequal to ψ = 25◦ and β = 30◦. The resulting stiffness Kspring required can be calculated according tothe moment equilibrium equation 3.6.

M

hservo· hstrut = Fspringhspring cos (γ) (3.6)

The stiffness required is dependent on the torque applied on the servo M , the angle γ between the springaxis and vertical position, the spring arm hspring about the center of rotation and the maximum availablespring travel sspring. Applying the equation 3.5 to the moment equation 3.6, yields the equation 3.7, thatallows to select the extension springs with a feasible combination of stiffness and stroke.

Ks = Mtan (ψ)tan (β) · 1

hspring cos (γ) · sspring(3.7)

3.3.5 Testing of the nose gear

The nose gear strut, already drop tested as discussed in subsection 3.3.2, has been characterized in termsof structural properties, through static tests with increasing weight between zero and the maximum valuethat corresponds to the bottoming out of the internal spring system.

The tests have been performed using a test rig mounted on the optical table provided in the shopof the Center for Aerospace Research, as shown on the left of figure 3.25. The loads on the strut havebeen applied with a screw mechanism and regulated by using a digital force gauge, while the resultingdisplacements on the strut have been registered using a digital caliper.

The stiffness of the nose gear strut is essentially carried by the internal spring system, regulated witha double acting spring in order to perform a progressive compression, as shown in the graph load/defor-mation shown on the right of figure 3.25.

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Figure 3.25. Static test for the characterization of the nose gear strut: the test rig is represented on theleft, the resulting load/deformation curve on the right.

As expected, the stiffness property of the nose gear is represented by two different linearized slopes,corresponding to the action of a single spring or double springs. The spring lengths have been selectedin order to attribute the load on the first spring until the acceleration reaches values of three times thegravity. As a consequence the nose gear can have a soft performance, to smooth the vibrations inducedby taxiing over unprepared runways, and provide additional stiffness during hard impact landing. Theactual stiffness coefficients of the nose gear for the two different soft and hard behaviour, are summarizedin table 3.10.

Table 3.10. Linearized stiffness coefficients for the nose strut during soft and hard performances

Soft performance Hard performanceKN 3.9 N/mm 6.5 N/mm

3.4 Integration of the landing gear with FTV7The present section illustrates the process of integration of the landing gear subsystems with two airframesof FTV7, respectively built for tailed and no-tail configurations. Both nose gear and main gear systemsare attached to the flat belly of the aircraft through specific structural components inside the front andmiddle bay of the airframe.

3.4.1 Integration of the main gearMain Gear location

The main landing gear has been placed at different locations with respect to the center of gravity position,as already discussed in subsection 3.3.2. Every location of the main gear meets the tip-back and take-offrotation requirements and ensure a load distribution of 90% on the main gear and 10% on the nose gear.The skin of the airframe has been provided with slots allowing an easy change of the main gear positionbefore each flight test, whenever required. The main landing gear positions with respect to the fuselagefront station (F.S.), depending on the center of gravity variation with respect to the reference meanaerodynamic chord length, are shown in table 3.11.

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Table 3.11. Main landing gear position with respect to the Fuselage Station

C.G. % w.r.t. MAC C.G. position [mm] Main gear position56% 749 800.1757% 755.66 807.5758% 762.32 814.9759% 768.98 822.3760% 775.64 829.7761% 782.30 837.1762% 788.96 844.1763% 795.62 851.9764% 802.28 859.37

Main landing gear attachment to the airframe

The flat part of the leaf has been designed in order to increase the capability of the strut to withstandthe stress concentration, induced by the bending moments that are developed by the vertical loads onthe wheels. In fact, the bending capabilities of a structure increase if the connection points are distanced,since the resulting stress distribution is more uniform, as demonstrated from the structural simulationsperformed and discussed in subsection 3.2.4. Six attachment points for bolts or rivet nuts connectionsare used in a three plate junction, that includes the leaf strut, the sandwich skin of the airframe and twointernal attachment plates inside the middle bay of the aircraft, shown on the right of figure 3.26. Theinternal plates, in Aluminum 2024 with 1/16" thickness, are used not to overload the skin of the aircraft.Their function is fulfilled where the washers and bolts are acting and so a symmetric shaped cut-out hasbeen designed to save weight and reduce possible interferences with internal components.

The final design of the flat surface includes three buttonholes in the areas where the material is lowstressed in order to reduce the weight of the landing gear and also facilitate the integration with theairframe and the electronic speed controllers (ESC), shown on the left of figure 3.26. The latter cannotbe moved from their initial location because they need to be ventilated for all the duration of the flight,and for this reason they have been mounted over the main leaf strut.

The damping capabilities to attenuate impact loads on main landing gear, have been improved thanksto rubber cushions, 60A in the durometer scale, located in between all the structural plates.

Figure 3.26. External and internal integration of the main landing gear with the airframe

3.4.2 Integration of the nose gearThe nose gear assembly is fixed to the flat belly of the airframe, through the same bolt holes used forthe launcher hook catapulting system of the aircraft without undercarriages. The first step has beento smooth the flat belly with sand paper P100 and sterilize the surface with alcool, in order to ensurea mating as large as possible. Then the carbon fiber plate has been applied and the attachment platein Aluminum 7075 has been bonded on, using epoxy adhesive EA-608. After a full curing time of 24hours, the attachment plate has been bolted in the four screw holes and all the assembly has finally beenmounted, as shown on the left of figure 3.27.

The same epoxy has been used to attach the servo 3D support to the inclined front wall of the nosebay. The nose gear cylinder passes through the skin of the aircraft and is located between the battery

28


packs that need to be there for trim and balance purpose. The final integration of the nose gear with theinternal front bay is shown on the right of figure 3.27.

Figure 3.27. Nose landing gear external and internal integration with the airframe

3.4.3 Belly Pan design

The landing gear parts, including all the fasteners used to attach the component to the airframe, havebeen covered by a foam belly pan that has basically two functions: streamline the flow on the bottompart of the fuselage and provide additional safety in case of emergency belly landing consequent to afailure of the landing gear. The foam belly pan has been provided with sufficient internal room to allowthe shifting of the main landing gear and not to interfere with the steering system mechanism.

3.5 Weight, cost and conclusions

The weight and cost distributions, relative to one manufactured landing gear set for FTV7, are shown inthe charts in figure 3.28. As already largely discussed, the main goal of the landing gear design, has beento keep as low as possible the weight and cost increasing respect to the original budget of the aircraft.

The weight distribution results from the mass evaluation of all the components used for the landinggear. The total weight of the two undercarriages results 5% of the maximum take-off weight for the tailedconfiguration of FTV7, in line with the general mass distribution of landing gear for comparable aircraft.[30] The weight is distributed as 74% for the main gear, mostly due to the beefy leaf strut, and 26% forthe nose gear, in majority represented by the nose gear strut and wheel.

In order not to highly affect the mass properties and inertias of the aircraft, a weight reduction analysisthat considers all the bill of materials and components used into the aircraft, has been performed inconjunction with the other responsible CfAR teams of the aircraft. The most candidate components tobe revised in order to allow an integration of the landing gear, without altering significantly the weight ofthe aircraft, has resulted to be relative to the batteries and parts used for the previous catapult take-offlaunches. The reduced capacity of the batteries and the removal of all the components correlated to thecatapult launch system, has allowed to integrate the landing gear without change the previous MTOWestimations.

The cost distribution results from an economic analysis using the PO pipeline orders in reference[31]. The cost of the off-the-shelf components is relatively low in comparison with the production ofcustom components, even though extra time has been required to edit the standard parts and facilitatethe integration with the overall landing gear design.

29


Figure 3.28. Weight and Cost distributions for the production of one landing gear set for FTV7

Conclusions

The aircraft FTV7 in both tailed and tailess configurations, with undercarriages installed, are now readyto take-off and land for the first time on a dedicated airstrip. Before a flight test campaign, it is necessaryto prepare all the settings and parameters for the autopilot, which is expected to have control of theaircraft from the initial taxiing until landing. The next chapter will introduce all the required groundmock-ups to prepare the aircraft with undercarriages for a flight test campaign, including the planningof all the activities connected to the landing gear analysis during the most critical phases of flight.

30

Chapter 4

Ground mock-ups and flight testplanning for FTV7%

Once the undercarriages have been integrated with the airframe, the FTVs are ready to be subjectedto ground mock-ups in order to estimate the expected control parameters to tune the autopilot. Thefirst part of the present chapter describes all the necessary steps to prepare the aircraft for ground andflight tests, with full commands imposed by the autopilot. Then the ground testing (taxiing and take-offrun), performed in the airstrip where the aircraft is suppose to take-off for the flight tests, is presentedmentioning to the necessary improvements suggested by the conditions of the runway. The last section ofthe chapter describes the flight test plan and introduces the process of flight test data analysis using theFlight Data Post-Processing software developed by Sean Bazzocchi and Jenner Richards at the Centerfor Aerospace Research.

4.1 Preparation of the aircraft for ground and flight testingSince all the phases of ground and flight testing are supposed to be controlled by the autopilot, specificground mock-ups are required in order to estimate mass and inertia properties of the aircraft and thenecessary control parameters for take-off run and steering.

4.1.1 Ground mock-upsCenter of gravity trimming

The first operation is to trim the aircraft for each position of the main landing gear and so for centerof gravity in the range 56% ÷ 64% of the mean aerodynamic chord. The trimming has been done withdifferent disposition of the movable components inside the front and middle bay of the airframe. Totest the effective mass and balance at the specific center of gravity position, the aircraft has been liftedand pivoted about a steel rod support attached to the middle bay at the corresponding c.g. position.The procedure has been done for each landing gear position and for the two configurations with tail andtailess, with weight respectively equal to 13.6 kg and 12.9 kg, including undercarriages.

Bifilar pendulum test

The Bifilar pendulum test is an approximate inertia calculator that uses a ground test rig and interactionamong different softwares including data collecting and processing to estimate the inertias of the aircraftin the three representative directions (yaw, pitch, roll). The procedure can be accurate only if all thecomponents, including the landing gear, are right in place as established from the center of gravitytrimming.

The aircraft is attached to a specific jig that is pivoted to different hooks, depending on the inertiaproperty to test (pitch, yaw and roll configuration).

The sinusoidal input is tuned on the autopilot with a specific frequency and the response collectedon the inertia measurement unit is transferred to the VectorNav software that creates the correspondinggraphs. The set-up data for the BFP test, including the aircraft assembly and jig mass, the jig fiber’sspacing and length and notes for the configuration tested, are collected in an Excel BFP CommissioningCard, ready to be used for post-processing.

31


The resulting inertias are calculated using the post-processor MatLab BFP software developed byJenner Richards at the Center for Aerospace Research. The software needs inertia inputs from theBFP Jig SolidWorks model calculated about the experimental coordinate system, that is the center ofgravity coordinate system of the full assembly, aligned with the CfAR aircraft coordinate system andlocated at the front fuselage station. The inertias of the aircraft are found removing the inertia of thejig about the experimental coordinate system from the total inertia of the assembly aircraft/jig. Thelatter is calculated, about the same point, from inputs of the experimental pendulum test according tothe relation expressed in equation 4.1, where m is the overall mass, D is the distance between the jigfibers, h is the length of the fibers and T is the period of the input signal. [32]

I = mgD2T 2

16hπ2 (4.1)

The moment of inertias about the front reference system are then calculated applying the parallel axistheorem. Note that the inertias for FTV7 have been calculated with landing gear located at 56%, whilethe inertia for the other configurations can be calculated applying the parallel axis theorem, knowing theinertia of the main landing gear from SolidWorks mass properties.

Mapping of the servo steering system

The ground maneuverability of the aircraft, controlled directly from the autopilot, is guaranteed by acareful mapping of the electric servo dedicated for the steering system. The operation has been accom-plished using a goniometer to check the angular position of the nose gear respect to the neutral positionand sending specific pulses from the ground station that is connected to the autopilot.

It has been decided to use a total steering capability of 28◦, due to the large width of the airstrip andso the not required classical steering capability of bigger aircraft operating in airports. The next stephas been to find the autopilot commands to maximize and minimize the steering ability selected. Thecorresponding pulse has been divided in ten mapping points and the resulting pulse spacings, measuredin µs, have been tuned in the autopilot command in order to register the effective angular position ofthe nose gear. The data of the servo mapping are collected in the table and graph shown in figure 4.1.From the resulting control mapping it is possible to notice that the response of the steering system tothe autopilot pulse input is almost linear in the selected range of steering ability.

Figure 4.1. Tuning of the steering system control parameters for the autopilot

Mapping of the thrust control parameters

The thrust control parameters for the autopilot have been mapped through a thrust test mock-up. Theaircraft with undercarriages has been placed on ground and connected through a rope to a load cell. Thethrottle command has been mapped with respect to the throttle percentage needed for all the phase oftaxiing, take-off and flight, following a similar approach described for the steering servo mapping. Theminimum (0% throttle) and maximum thrust (100% throttle) defined by the propulsion system, have been

32

4 – Ground mock-ups and flight test planning for FTV7%

used to read the corresponding values of pulses on the autopilot, then the command has been equallyspaced in ten mapping points and used as input in the autopilot. Finally the thrust mapping has beencompleted reading the corresponding values of thrust on the load cell.

4.1.2 Landing gear toolkitAll the landing gear spare part hardwares have been organized in a specific toolkit with the relativechecklist annexed on the top, in order to facilitate the pre-flight preparation and the necessity to quicklyrepair or replace the components in case of relevant damages. The toolkit contains even all the landinggear assembly drawings with the description of the component list and relative part number. A sampleof drawing, relative to the nose gear strut, is shown in figure 4.2. All the other mounting drawings of thelanding gear for FTV7 are illustrated in Appendix B.

Figure 4.2. Component list and mounting drawing for the Nose Gear strut assembly 1

4.2 Ground testingThe FTV2B and FTV2C with undercarriages have been ground tested on the runway where the aircraftis supposed to take-off and land, in Merrit - Douglas Lake (Canada). The goal of the tests has been toverify the performance of the aircraft with landing gear in all the phases that precede the flight, with theintention to make improvements if necessary.

4.2.1 Taxiing of the FTVsDuring the preliminary taxiing tests, the aircraft has resulted not to respond as expected from theautopilot nose gear inputs. Manual testing has revealed compromised yaw authority due to significantplay in the coupling between the steering shaft and linkage horn, probably due to the quite uneven runwayconditions. As temporary solution it has been decided to add a drop of cyano-acrylate glue in order toreinforce the clamp fit. The aircraft has so regained yaw authority and several taxiing tests have beenperformed in autopilot and manual mode.

33


4.2.2 Take off run testing of FTV 2BThe take-off run tests are the next step towards a fully operating aircraft, ready to be flight tested. Theaircraft has been attempted to be controlled in the centerline of the airstrip in a take-off run sequence.

The initial response has showed out of control of the aircraft when high speeds were achieved, dueto the airstrip crack induced bouncing. Tuning PID loops on the autopilot control have been necessaryto try and get the aircraft running true in the take-off run even if the runway conditions were not asexpected. After several attempts it has been agreed that larger wheels on the main gear could havehelped to let the aircraft controllable with reasonable values of PID parameters.

4.2.3 Improvements on the landing gearThe problems encountered in taxiing and take-off run testing, due to the bumpy conditions of the airstrip,have led to the necessity of making improvements in the design of some components of the lading gear.

Due to the steering fit slopping, it has been decided to improve the design of the horns used in thesteering system, increasing the contact area in between the servo shaft and the horn. In addition thetwo horns have been provided with a full open clamping slot, in order to increase the effectiveness of thefitting with an higher torsion clamp.

The difficulties encountered during take-off run over an high irregular runway, have indicated thenecessity to swap for bigger size of wheel assemblies for the main landing gear. Tires with 25% biggerdiameter respect to the previous solution adopted, shown in figure 3.3, have been selected. The aircraftattitude has been maintained the same of the previous design, compensating the increasing of the height(12.5%) on the main gear contact point by adopting a nose gear with the original strut length (as frommanufacturing), previously described in subsection 3.3.2.

4.3 Flight test planningThe FTV2B and FTV2C aircrafts, with the improvements on the landing gear adopted from the groundtesting results, are now ready to be flight tested even in a rough and bumpy airstrip, like the one inMerrit - Douglas Lake (Canada). The new inertias of the aircraft with a new set of gears and nose strut,can be estimated removing the contribution of the previous components and adding the inertia of thenew components to the aircraft inertia about the front fuselage station, as described in subsection 4.1.1.

The planning of the flight tests, in particular regarding the implications on the landing gear, isdescribed through the following steps:

1. Full take-off operation of the wheeled FTV2B and FTV2C in autopilot mode, with ground stationofficer (GSO) ready to control the aircraft if something does not go as expected.

2. Flight of the aircrafts in autopilot mode, with a flight path based on the endurance and flightrequirements agreed with the stakeholders.

3. Landing of the aircrafts on the same airstrip used for take-off.

4. Check the aircraft conditions after the full operational flight.

5. Analyze all the critical flight test phases for the landing gear using the Flight Data Post-Processingsoftware.

6. Use all the processed data to redesign components, if necessary, and to make decisions on thelanding gear design for the larger flight test demonstrator 16.5%.

4.3.1 Flight test data processingThe response of the aircraft with landing gear during all the flight phases of interests (take-off and landing)can be evaluated using the Flight test data processing software provided at the Center for Aerospaceresearch and developed by Sean Bazzocchi and Jenner Richards. The general graphical interface of thesoftware, for a typical landing overview of a previous flight test of FTV2B without landing gear, is shownin figure 4.3.

34

4 – Ground mock-ups and flight test planning for FTV7%

Figure 4.3. Landing overview using the Flight test data processing software

The landing gear designer can obtain all the telemetry data of the aircraft during take-off and landing,including the variation of altitude, vertical rate of descent, take-off and landing horizontal speed, attitudeof the aircraft in terms of roll, pitch and yaw, and even information regarding the accelerations developedon take-off (horizontal acceleration ax) and landing (vertical acceleration az) and so the effective load onthe wheels developed at the ground contact. Specific points on the flight phase (for example the instantof touch-down or take-off) can be evaluated using a data cursor extrapolation tool. All these data can bevisualized on multiple plots: for instance, the center body of the software interface shown in figure 4.3 isrepresenting the trend of the altitude and true air speed in a landing of FTV2B without undercarriages,and the g developed in all the directions.

The analysis of flight test data is followed by the development of flight test statistics for importantdesign parameter, such as the g developed on landing and the rate of descent of the aircraft beforetouching-down, that can be compared with the values used for the design.

The data can be analyzed for each configuration of the main landing gear and for the determinationof the optimal center of gravity location for take-off and landing, from a landing gear point of view, thatcan be useful to guide the design of the landing gear for a larger flight test vehicle demonstrator. Theobtained time-history plots can be even useful to prove the reliability of design structural simulationsand prove the development of a Simulink model of the aircraft response with undercarriages.

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36

Chapter 5

Landing Gear design for FTV16.5%

The design of the new generation blended wing body aircraft has been started for a bigger 16.5% FTVscale, in parallel with the 7%, in the process of scaling towards to a flight test vehicle that is morerepresentative of the full scale aircraft. The large scale demonstrator is expected to represent as closeas possible the same behavior of the full-scale in all the flight phases, including take-off and landingoperations that require the design of a reliable undercarriage system. The landing gear system foran aircraft with size and weight of FTV 16.5 is not readily available form radio commanded aircraft orultralight applications, so it has to be designed and produced from scratch. The main goal of the designerof the landing gear for big UAV scales, is to search at the same time for a simple landing gear in termsof weight and cost, and a complex machine capable to support and transfer the largest local loads fromthe ground to the airframe.

This chapter highlights all the progress done for the design of the landing gear towards to the preli-minary definition of the working laws involved in its functionality. An initial brief presentation of all theaspect concerned in the design of the fully operative system, leads to the determination of all the basicfeatures of the undercarriages for FTV16.5, before entering in detail of the conceptual design, followedby the selection of the most suitable concepts and preliminary design, of both main gear and nose gearsystem with the relative mechanical subsystems.

5.1 Design process for the 16.5% Landing Gear systemThe design of the landing gear system for a new big scale aircraft is a self-contained project that passesthrough all the typical design phases, but at the same time it requires close collaboration among allthe different design team, including airframe, sizing, recovery system, propulsion, flight control andaerodynamics. The work break down structure (WBS), that defines the guideline framework for thedesign progress of the landing gear system for FTV16.5, is shown in figure 5.1.

The design of the landing gear follows the same schedule and deliverables time of the airframe designand is accompanied with all the necessary documentation and reports, that represent the basis of discus-sion with the stakeholders. The present thesis work covers the first three phases of the project, performedin the period in between May and September 2018.

The first approach has been to define the project charter that tracks all the function required for thefinal product and some basic requirements, agreed with the stakeholder, in order to guide all the designprocess. In addition to the all aspect involved in the design of the landing gear for FTV7, discussed inchapter 3, the design for a bigger scale UAV requires the development of a suspension system, dedicatedfor the shock load management, and braking system to stop the aircraft within the length of the runway.The project is also focused on the development of specific Test rigs that allow the testing of the integrityand functionality of the designed landing gear. Specific tests are required to verify the response of shockabsorbers, the speed rating of the tires, the energy dissipation upon braking and the structural responseof the aircraft on the landing gear during static and drop tests. [4]

The design starts with the evaluation of all the possibilities and concepts for the main systems thatdefine the basic operation of the undercarriage and then moves, through a market analysis and analyticalhierarchical prioritization of the proposed solutions, to the selection of concepts. The successive phaseregards the definition of preliminary design of the basic components needed, including the evaluation ofthe basics of braking and steering system. Off-the-shelf components are evaluated, adjusted and fullyintegrated into the design. At this level a preliminary CAD model is prepared and it defines the boundary

37


space needed for the following detailed sizing.

Once all the components needed have been characterized, the design moves to the detailed definitionof each system and subsystem, including the clarification of all the test rig procedures needed to certifythe design. At this point detailed structural and aerodynamic finite element based analyses are requiredto optimize the modeled components. The final objective is to generate the complete component list andassembly drawings that lead to the manufacture process and integration within the airframe.

Figure 5.1. Work break-down structure of the landing gear design

The undercarriage design is a multidisciplinary process that requires the use and interaction of differentsoftwares, as illustrated in figure 5.2. The basic features of the landing gear are defined using Excelspreadsheets and Matlab sizing tools. The next step is to translate the initial layout of the landing gearin solid CAD modeling. Once the design is refined, structural and aerodynamic simulation tools can beused to evaluate the responses of the landing gear to simulated loading cases that represent some specificphases of the flight. The design is then optimized and evaluated in the contest of the full aircraft, byusing Simulink models, ground testing and integration checks.

38

5 – Landing Gear design for FTV16.5%

Figure 5.2. Multidisciplinary process of the landing gear design

5.2 Basic features of the 16.5% Landing Gear

The basic characteristics of the landing gear for FTV16.5 have been determined after a scaling processfrom the 7% and 100%, guidelines from FAR 23 regulations and typical design steps described in reference[10] and already implemented for the design of the landing gear for the 7% aircraft. The landing gear forthe large scale demonstrator is assumed to have the same configuration (tricycle and fixed) and geometricsimilarities with respect to the 7% landing gear and the design load factor used to size the structuralelements is decided to be the same suggested by certification of landing gear for FAR 23 regulations, asexpressed in subsection 2.2.1.

The first step of the design process has been to identify and list all the design requirements in an Excelspreadsheet file. The requirements, as already done for the FTV7, are divided by category depending onthe type of the feature they affect: aircraft and performance, geometry, main and nose gear peculiarities.The list can be filtered by the origin of the requirement (scaling of Bombardier aircrafts, sizing designteam, airframe design team, parachute design team, FAR and MIL specifications, design book guidelines),status of the requirement (open, completed, deferred, no longer required), priority (high, medium, low)and date. An example of the Excel format sheet for the basic requirements is shown in figure 5.3.

The next stage of the design process has been to import all the requirements from Excel to MatLabin order to calculate the basic features of the landing gear, using the developed Matlab® parametercalculator software, described in Appendix A and already used for the FTV7 landing gear. At this pointthe design of the landing gear becomes iterative and all the changes in the landing gear variables shouldsatisfy the listed core requirements.

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Figure 5.3. Basic requirements for the 16.5% Landing Gear design

The next subsections describe all the basic aspects that have been considered to characterize theinitial layout of the landing gear for FTV16.5.

5.2.1 Landing Gear position and aircraft center of gravityThe importance to relate the landing gear design to the center of gravity envelope is to make sure, in aparametric approach, that the major geometric landing gear variables satisfy all the requirements. Thedesign approach in this case has been to scale and fix the nose gear position and determine from theC.G. envelope, contained in reference [2], the position of the main gear that better implies the respectof all the design requirements. The final position of the main landing gear has been determined after aseries of iterations and revision of all the landing gear parameters. In a tricycle configuration the optimallocation of the main gear relative to the forward position of the C.G. is governed by the take-off rotationrequirement. The consequent tip-back angle, that is the angle between the farther and lowest point of thefuselage or tail and the ground, should respect the tip-back angle requirement that prevents the aircraftfrom tipping back on its tail during rotation, as already introduced in section 2.1.

Take-off rotation requirement

The distance between the main gear and the most forward position of the c.g. is defined by the take-offrotation requirement, that allows the aircraft to rotate around the main gear in order to achieve the angleof attack required for take-off. An important parameter that defines the take-off rotation requirement isthe pitch acceleration θ, when the aircraft begins to rotate. The rate of change of the pitch rotationalspeed depends upon several parameters including tail geometric parameters, elevator control power,aircraft weight, rotational speed and distance between the main gear and the aircraft c.g. A typicaltake-off pitch angular acceleration for the FTV16.5 high maneuverable aircraft is θTO = 15[deg/s2]. [10]During the rotation the speed VR of the aircraft can be assumed to be a function of the stall speed Vs,as expressed in equation 5.1.

VR = 1.1 ÷ 1.3VS (5.1)

The forces and moments involved during take-off rotation are represented in the body force diagramin figure 5.4. The represented vectors refer to the aerodynamic forces at the aerodynamic center of thewing and tail (if present), propulsion thrust, aircraft mass forces and friction force at the main gearcontact point. The equilibrium state represented refers to the instant following the nose gear lift-off fromthe ground, so no forces on the nose landing gear are considered.

40


Figure 5.4. Equilibrium about the main gear contact point at take-off rotation

The maximum main gear position depending on the take-off rotation requirement can be calculatedapplying a moment equivalence about the MG contact point, as shown in equation 5.2, where the pitchinginertia (Iyy−mg) about the main gear contact point can be calculated with the parallel axis theorem fromthe inertia of the aircraft about center of gravity (Iyy), as in the equation 5.3.

xm =Iyymg

θ −Wxcg −mazcg + Lxac−w +M0w − Lhxac−h −D · zD + T · zTL− Lh −W

(5.2)

Iyymg = Iyy + W

g(xm − xc.g.)2 (5.3)

Tip-back requirement

The tip-back angle αtb is the maximum aircraft nose-up altitude with the tail touching the ground andthe gear in fully extended position. It should be at least 5◦ larger than the take-off rotation angle andneeds to be equal or less than the angle A, measured from the vertical at the main gear location to theaircraft most aft center of gravity, as expressed in equation 5.4. [7]{

αtb ≥ αTO + 5◦

A ≥ αtb(5.4)

The A and αtb can be calculated according to equations 5.5 and 5.6, where zcg considered is the ver-tical position of the c.g. in the worst case scenario (higher position), xfp and zfp are respectively thelongitudinal and vertical location of the farther and lowest point in the worst case scenario (no tailconfiguration).

A = tan−1(xm − xc.g.

zg

)(5.5)

αtb = tan−1(

zfpxfp − xm

)(5.6)

Position of the main gear

The longitudinal C.G. envelope of the aircraft is quite wide (from 56% − 66% of the mean aerodynamicchord) because, as already discussed for the 7% aircraft, the new generation flight demonstrator needsto be tested with a movable mass and C.G. inside the aircraft, allowing the flight in both stable andunstable configurations. If the position of the main gear is wanted to be fixed, it needs to be located

41


at the forward minimum location that results from the aft location of the C.G (66%). This positionrespects the tip-back requirement, expressed in equation 5.4, but the resulting angles for the forwardlimit of the C.G. would be too much different, respectively A=34◦ and αtb = 18.6◦, and this implies anhuge amount of load needed on the elevon to trim the aircraft and allow the take-off operation, that isnot easy manageable in a no tail configuration.

The decision made at this point has been to privilege the design of a "semi-movable" landing gear,which ensure similar load distribution between the nose and main gear along the C.G. envelope. Moreoverthis configuration implies at the same time the easier take-off rotation (A and αtb angle similar), eventhough it adds complexity and extra structural components respect to the classic fixed gear. The mainlanding gear strut leg can be fixed before each flight test choosing from 4 position inside a swivel com-ponent, directly attached to the pivot point on the fuselage. The basic concept of the swivel componentwhich allows different settings of the main gear is shown in figure 5.5.

Figure 5.5. Sketch of the swivel component, necessary to rotate the main gear position

As the landing gear configuration changes, with a rotation of the strut leg about the vertical axis onthe pivot point of the component, the resulting wheel track dimension changes, maintaining the sameheight of landing gear. The variation of the track between the two main gear assemblies will be discussedin subsection 5.2.3.

The main landing gear different configurations, depending on the C.G. envelope that results in stableor instable positions, involve the main gear contact points,1 shown in table 5.1.

Table 5.1. Set of landing gear configuration for different position of the c.g.

Configuration Aircraft stability C.G. envelope Contact point [mm]A Yes 56 − 58% 1940B Yes 59 − 61% 1987.1C No 62 − 64% 2034.2D No 65 − 67% 2093

5.2.2 Landing Gear heightThe landing gear height is defined as the distance between the ground contact point and a reference inthe aircraft, in this case the front fuselage water line. As done for FTV 7, the initial approach has beento scale the dimensions of the full scale aircraft with the landing gear in fully extended, static and fullycompressed positions and its determination needs to be validated after checking the take off-rotationground clearance criteria expressed in section 2.1. The clearance angle αC is calculated according to

1The dimension of the main gear contact point are relative to the front fuselage station (F.S.)

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equation 2.4 and is proved to be less than αTO. The height and take-off ground clearance requirementimply a ground clearance of Hc = 95.80mm, as shown in figure 5.6.

Figure 5.6. Take-off rotation ground clearance requirement

5.2.3 Track and Turnover angleThe wheel track T is defined as the distance between the most right and most left main landing gear onthe ground, in a frontal plane view. Its determination for the FTV16.5 aircraft has been finalized afterscaling considerations and the necessity to move the main landing gear for different position of the centerof gravity. The minimum allowable value for the wheel track must satisfy the turnover angle requirements(turning controllability and ground stability), whilst the maximum position should meet the structuralintegrity requirement.

The turnover angle φ is determined from a drawing sketch procedure as shown on the left of figure5.7, referred to the scaled down track.

Figure 5.7. Geometric parameters involved in the determination of the turnover angle

Turning controllability on ground

The wheel track T plays an important role in the ground controllability of the aircraft and its determina-tion passes through the moment equilibrium about the gear contact, as illustrated in equation 5.7, that

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refers to figure 5.7.

Fc · zcg = W · T2 (5.7)

The turnover prevention requirement imposes that the moment generated by the aircraft weightW shouldbe higher than the one caused by the centrifugal force Fc. The consequent minimum track dimensionand turnover angle are expressed in the system of equations 5.8, where Rturn is calculated according toequation 3.4 stated in chapter 3.T >

2Fc·zc.g.

mg = 2V 2zc.g.

gRturn

φ > tan−1(Fc

mg

)= tan−1

(V 2

gRturn

)= tan−1

(T

2zc.g.

) (5.8)

Ground stability

The other case that can roll over the aircraft is a cross-wind taxiing, assuming in the worst case scenariothat the wind is perpendicular to the centerline of the aircraft, ass represented on the right of figure5.7. To prevent the aircraft from overturning during cross-wind conditions, the moment generated bythe weight about the main gear contact point should be higher than the one generated by the cross windforce. The requirement is critical for the lowest mass of the aircraft (zero fuel) and it is shown in equation5.9

T >2FW · zc.g.

mg(5.9)

The aerodynamic cross wind force FW is estimated according to equation 5.10.

FW = 12ρV

2wASCDS

(5.10)

The aerodynamic parameters that determine the cross-wind force are the wind speed VW (assumed tobe 50knots), the aircraft cross side area AS invested by the cross-wind, and aircraft side drag coefficientCDS

(the value is in between 0.3 and 0.8). [10]

Structural integrity requirement

The maximum value of the wheel track is determined from a structural point of view that interests thedeflection of the strut leg, attached to the pivot point on the fuselage. In this analysis it is assumed thatthere is no suspension system and the strut leg is directly loaded when the ground force is applied onthe wheel. The most critical condition is achieved when the static ground reaction Fmmax , correspondingto the aft limit position of the c.g., is applied on the wheel and the deflection is the maximum value ofwheel travel with 2 inches of clearance to the wing. [33]

The corresponding maximum track can be estimated by using equation 5.11, that comes from theapproximation of the strut leg as a beam, applying the displacement equation 3.3 stated in chapter 3.

Tmax = 2

√[3EIeq · ymaxPmmax

] 13

+ y2max (5.11)

Final determination of the wheel track and turnover angle

The final wheel track of the main landing gear is expected to change among the four configurationsdescribed in subsection 5.2.1, as the main strut rotates about the vertical axis on the pivot point. Itis assumed that configuration B (at 60% location of the center of gravity) has the scaled down trackdimension from the full scale aircraft, shown in figure 5.7, in order to keep as minimal as possible theresulting variation of the track dimension for the other configurations. The resulting track and turnoverangle for each configuration, shown in table 5.2, respect the limit values stated by the turnover angleand structural integrity requirements. As observed, the track offset respect to the scaled-down nominaldimension is quite low:+1% for configuration A, -2% for configuration C and -4.3% for configuration D.

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Table 5.2. Track and turnover angle for each main gear configuration

Configuration Wheelbase Rturn Tmin Tmax T φA 1292 2582 750 1470 1344 40.76B 1339 2676 724 1520 1337 41.80C 1386 2770 699 1570 1323 42.65D 1445 2888 670 1637 1293 43.83

5.2.4 Design loading conditions for the nose and main gearThe loading distribution between main and nose gear is largely affected by the wheelbase, defined as thedistance between the nose and main landing gear and in this case it is determined by the fixed scaled-down position of the nose gear and the resulting position of the main gear for each of the 4 configurationsdescribed in subsection 5.2.1. The static load distribution can be determined from the arm ratio of thenose and main landing gear with respect to the most critical position of the center of gravity, that isrespectively the forward limit for the nose gear loading and the aft limit for the main gear, as expressed inequation 2.1 in section 2.3. The resulting combination of loads for all the main landing gear configurationsand for the most critical case of C.G. for each configuration are summarized in table 5.3.

Table 5.3. Static loading cases on the nose and main gear for each landing gear configuration

Configuration Wheelbase [mm] C.G. shifting PNmax [kg] PMmax [kg] Load RatioA 1292 56-57-58 % 16.8 77.83 9.6%- 89%B 1339 59-60-61 % 16.86 78.17 9.6%-89.3%C 1386 62-63-64 % 17 78.50 9.7%-89.7%D 1445 65-66-67 % 16.94 78.14 9.7%-89.3%

The vertical loads on the nose and main gear are increased during braking and take-off, accordingto the equations 2.2 and 2.3 stated in section 2.1. In this case, the highest value of the vertical C.G.envelope of the aircraft is used because it represents the most critical situation with higher load. Theresulting maximum static and dynamic loadings for the nose and main gear, summarized in table 5.4, areutilized for the wheel and tire design as well as for the shock absorber selection.

Table 5.4. Total loading cases on the nose and main gear for each landing gear configuration

Configuration PNdyn [kg] PNTOT [kg] PMdyn [kg] PMtot [kg]A 22.23 39.03 25.40 103.23B 21.45 38.31 24.51 102.68C 20.73 37.73 23.69 102.19D 19.88 36.82 22.72 100.86

5.2.5 Impact loading condition for the landing gearThe landing cases that the designer should consider are nominal landing and parachute emergency landing.The design of impact behaviour on landing for the specific case of FTV16.5 can be outlined followingeither a scaling procedure from the full scale aircraft or applying the typical design procedure describedin references [7] or [8].

Scaling procedure

The input for the scaling procedure is to assume that the performance of the landing gear for FTV16.5is the same in terms of vertical deflection of the full scale model. The second parameter essential to

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determine the required impact acceleration on landing is the vertical rate of descent, calculated in thecase of nominal landing with maximum take-off weight according to the equation 2.5, introduced inchapter 2.3 and valid for aircraft certified according to FAR 23. In case of parachute emergency landing,the typical rate of descent is selected in a trade-off with the parachute design team.

The resulting average impact load factor, developed on landing, can be calculated as shown in equation5.12, obtained simply applying the basics of kinematics and dynamics principles.

NG = V 2z

2 · d · g(5.12)

The resulting normalized accelerations developed on landing for the two design case scenarios is shownin table 5.5.

Table 5.5. Scaling design procedure: impact accelerations developed on nominal and parachute landings

Compression [d] Sink speed[Vz] Impact accelerationNominal 64.11 mm 7.33 fps 4gParachute 64.11 mm 15 fps 16.6g

Classic design procedure

The classic design procedure for the landing gear design, already used for the design of the undercarriagesfor FTV7, is to select from the beginning the design ground load factor and then calculate the requiredwheel travel to perform the consequent impact acceleration. A typical ground load factor for small aircraftcompliance with FAR 23 regulations is 3, as stated in the section 2.2. The wheel travel, corresponding tothe 3g nominal landing, is obtained using the equation 5.12 and results to be 25% more than the nominalscaled down compression.

In the emergency landing case, it is assumed that there is the bottoming out of the internal suspensionsystem and that the wheel travel is higher than the nominal value. This procedure allows to lower theimpact normalized acceleration on parachute landing from 16.6g, calculated according to the scalingprocedure, to 11.5g. The load factor for the two nominal and emergency cases, using the described designprocedure is represented in table 5.6.

Table 5.6. Classic design procedure: impact accelerations developed on nominal and parachute landings

Compression [d] Sink speed[Vz] Impact accelerationNominal 84.8 mm 7.33 fps 3gParachute 92.9 mm 15 fps 11.5g

Selected design procedure for the landing gear of FTV16.5

Due to the higher impact acceleration to dissipate in a scaling design procedure, it has been agreed withthe stakeholder to design the landing gear in order to have the wheel travel for nominal and parachutelanding that results in the lower impact acceleration.

The graphs in figure 5.8 show the trends of the impact average acceleration developed on landing infunction of the compression and rate of descent.

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Figure 5.8. Design procedures for impact loading condition: on the left the scaling design procedurewith influence of the parachute rate of descent; on the right the classical design procedure withinfluence of the design wheel travel on landing

The design impact acceleration is lower if the admitted rate of descent decreases or the compression ofthe landing gear increases. The classical design procedure has been chosen, since the rate of descent usedfor the design of the parachute system has been resulted hard to change significantly and also becausethe suspension system is expected to manage the impact loads with the maximum possible wheel traveland the minimum shock stroke.

5.3 Conceptual DesignIn the first conceptual phase of the project the designer is faced with a variety of possible configurationsfor the most important components of the main and nose gear and their attachments with the airframe.As done for FTV7, the general design guideline has been to use, where possible, off-the-shelf components(COTS) in order to reduce the complexity and cost of the design work. At this stage the landing geardesigner is aware of the ground tests needed to prove the structural integrity and functionality of theselected components and if necessary, redesign parts and adjustments.

5.3.1 Main gear conceptsThe present section summarizes all the concepts and the guidelines used for the conceptual design ofthe tricycle landing gear. The concepts are shown at level of configuration (external and internal) andcomponents needed, analyzing standard components whenever possible.

External layout configurations

The external layout of the main landing gear is guided by the type of attachment with the airframe. Thetwo primary options are the fuselage and the wing and the final choice influences the take-off and landingperformance as well as the ground stability.

The full scale aircraft has the main landing gear attached vertically to the wing: this configurationon FTV16.5 would imply an optimal shock load distribution inside the suspension system and a constantvalue of the wheel track between fully extended and fully compressed position of the gear, but on theother side, it would require additional structural components to transfer the loads from the wing ribs tothe bulkheads of the fuselage, that means additional load path complexity and weight.

A fuselage attachment represents a valid alternative since the fuselage is wide enough to providethe required wheel tracks and to carry directly the loads developed on the landing gear. Although theresulting load distribution in this configuration on the shock absorber is higher, it has been preferredbecause it can allow the variation of the ground contact point on the main gear, as discussed in subsection5.2.1.

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Internal layout configurations

The internal configuration of the main landing gear is essentially guided by the suspension system andstructural supports for the attachment on the bulkhead. Different configurations have been comparedand evaluated in terms of the resulting kinematics and required components, and the most interestingones are sketched in figure 5.9.

Figure 5.9. Sketches of the concepts for the internal layout configuration of the main landing gear.

The sketches show, for each configuration, the kinematic motion from fully extended to fully com-pressed position of the landing gear, the pivot points on the bulkhead, the wheel travel and the strokeon the shock absorber. The main features for each configuration are summarized as follows:

• Configuration 1: The first solution presents a pivot point on the fuselage frame with an internalshock absorber that is disposed with a certain angle with respect to the strut leg. The motion ratio,between the wheel travel and the stroke of the shock, is completely governed by the geometric armratio between the internal and external strut leg. This results in an efficient and simple kinematicswith the possibility to easily integrate a common attachment plate for the shock absorbers of thetwo main gears.

• Configuration 2: The second configuration is similar to the first one in terms of pivot point onthe fuselage frame, but it presents an additional component that can be used to increase the motionratio and so have the maximum wheel travel, with the minimum stroke on the shock. On the otherside this adds complexity to the design, due to the necessity to integrate the motion of the shockabsorber with a third linkage component and an additional hard-point on the fuselage.

• Configuration 3: The last solution presents an internal pivot point connected to the fuselagebulkhead, with an internal shock absorber disposed in the upper-left part. This results in a lowrequired internal space and easy integration with the internal layout of the fuselage. On the otherside, the available stroke is lower and cut-outs on the fuselage are needed to allow the strut leg torotate in loading conditions.

The choice of the most suitable configuration depends even on the type and size of the off-the-shelfcomponents needed for the suspension system. The final selection will be the first step to start thepreliminary design of the landing gear.

Components off the shelf (COTS)

As already largely discussed in chapter 2 and 3, the simplicity and cost of the design of the landing gearis guided by the availability of standard components from UAV and ultralight aircraft applications. Adetailed market research has been done for each basic component needed for the main landing gear: tires,wheels, brakes and shock absorbers.

The tire has essentially the function to support the aircraft structure off the ground, help to absorbthe shock loads depending on the inside pressure, transmit the accelerations from the ground to thelanding gear structure and braking forces to the runway surface, help to maintain or change the direction

48


of motion. The main parameters, that guide the selection of the proper tire, are the geometric outsidediameter and width, and the total static loads that the landing gear is expected to experience, statedin table 5.4. It has been established that the maximum geometric size2 of the tires should have notexceeded the 25% more of the scaled down dimensions from the full scale, in order to keep minimum thedifference of behavior, between the FTV and full scale aircraft, during the phases of take-off accelerationand ground effect on landing. [33]

The scaled down size from the full scale aircraft is 6"x2"-1.5" and such a type of tire has been difficultto find in off-the-shelf market, since it is in between radio commanded and ultralight aircraft applications.The loads that the tires should withstand are easily manageable, but the speed target for FTV16.5 isquite high for almost all the tire in that application range, only rated for low speed tail-wheel applications.These tires are applicable to the main gear concept only if they can be ground tested and certified for thespeed range of interest, otherwise the decision should go towards tires for other UAV applications (suchas the Sonex Jet and the Boeing experimental drone X-36, introduced in section 2.3), where the ratedspeed values are more comparable, even if they have bigger size. The main results from the tire marketresearch are summarized in figure 5.10.

Figure 5.10. Off-the-shelf tires from the UAV and ultralight aircraft market

The wheel size required for the main gear of FTV16.5 is guided by the bead seat diameter of theselected tire. Different solutions have been evaluated and the main feasible ones are illustrated in figure5.11. The classical solutions are made from aluminum forged alloy, usually in dural 2024 or ergal 7075and they are produced in two halves then joined together by a number of tiebolts. The hub is designedto house the wheel bearings, that are taper roller type and sealed to ensure their grease is not ejected athigh speed. Each of the solutions proposed are rated for high static capacity and load limit, with respectto the needs for applications to FTV16.5. An innovative solution is offered by the tubeless STS wheelwith integrated brakes machined in ergal 7075. A tubeless solution has essentially three advantages inthe aeronautical field: safety and less probability of punctures, lightness avoiding the use of a tube/airchamber, easy affordability. However the minimum tubeless off-the-shelf option found in the market hasa rime diameter of 4", as shown in figure 5.11, that does not allow the integration with a 6" tire. [34]

2The geometric size of tire is represented as D0 × W − DW , where D0 is the outside tire diameter, W is the widthand DW is the wheel diameter, or tire bead seat diameter.

49


Figure 5.11. Off-the-shelf wheels from the UAV and ultralight aircraft market

The brakes can be either off the shelf solutions from ultralight applications for the wheels describedbefore, or they need to be integrated from other applications, such as bikes or quad karts. The brakes canbe distinguished for type of actuation (pneumatic, electro-magnetic, hydraulic) and type of brake contact(integrated disk brake, external or internal caliper3, drum internal brakes). Most of all the applicationsfor aircraft in the range of FTV16.5 use hydraulic brakes with external or internal caliper, so the marketresearch has been addressed to this type. The feasible solutions for the wheels described in figure 5.11,are presented in figure 5.12. The smaller wheel solution can be integrated with disk brake and caliper forbicycle applications, as already implemented for the X-36 UAV aircraft, if no specific braking distancerequirement are needed and the length of the runway is much more higher than the scale of the aircraft.If the aircraft needs to be stopped within a certain distance and an higher stop kinetic energy is required,the applications should move towards calipers designed and sized for ultralight applications, in compliancewith FAR 23 regulations. In alternative, the STS tubeless wheel described, presents the possibility tobe integrated with a brake system that works as a clutch, with pistons that generate an axial force andthe disks packed alternatively with friction material and steel. This solution is extremely light weightand it results highly performing, because the friction material has a larger surface than a classic brakingsystem, and so the same braking torque is obtained with half hydraulic pressure. [34]

Figure 5.12. Off-the-shelf brake solutions for FTV16.5 applications

The internal suspension kinematics and structural response is mostly dictated by the shock absorber

3External or internal refer to the position of the caliper with respect to the wheel diameter.

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selection. In general, for fixed landing gear, a solid strut leaf or a rubber bungee is a suitable options, butin this case due to the scale of the aircraft and to the applications in not perfect smooth runways, it hasbeen decided to rely all the suspension properties (stiffness and damping) to apposite shock absorbers.

The first approach has been to search for components among bicycle/bike and ultralight aircraftapplications, in order to alleviate the impact cost of a custom solution. The main interesting solutionsfrom the preliminary market research are shown in figure 5.13. The simpler solutions is a coil springwith separate elements for the compression (spring) and damping (oil, piston and orifices), even thoughit needs a relative wide amount of space. A more compact solution is represented by an air spring withan internal chamber, design in order to use the air for both compression and damping properties. Theair spring shock absorber shown, has ten settings of rebound control and three position for compressionadjustment with apposite rotating valves. The best solutions in terms of rebound and compression isan oleo-pneumatic shock absorber with separate chambers and fluid for the two properties, in order tomaximize and optimize the fluid application for the specific task required. However, this is not an easysolution to integrate with small airplanes, especially if the shock absorber needs to be allocated insidethe aircraft, due to the large size of the shock and to the high cost and weight.

Figure 5.13. Off-the-shelf shock absorbers for FTV16.5 applications

Custom components

The off the shelf components described are expected to be integrated with a custom strut leg that canrotate about a swivel bracket before fixing the position for take-off and landing. The conceptual design ofthe strut has started from the evaluation of typical material that can potentially be used for the customcomponent. The geometry has been modeled considering the nominal outboard track at the position ofthe c.g. at 60% of the M.A.C. and following the classical procedure already used for the 7% and describedin subsection 3.2.3. The part, modeled in SolidWorks, is shown in figure 5.14 with all the typical designelements that include pivot points, internal and external arm, house for the wheel axle.

Figure 5.14. Conceptual geometry of the main gear strut leg

Different materials, typically used for landing gear strut components (Steel, Aluminum, Carbon fiber

51


laminate), and thickness have been used to evaluate the best conceptual application with the same overallgeometry. The structure has been modeled with TRIA solid mesh elements and the boundary conditionsapplied are hinge constraints for the two pivot points, that connect respectively the strut to the fuselageand to the shock absorber, and remote vertical load at the location of the wheel, corresponding to thedesign impact load. The material and thickness have been changed in an iterative process in order to keepthe structural safety factor equal to 1.5, commonly used for sizing the aircraft structure. The analysisresults, illustrated in figure 5.15, show that the best combination for a strut leg for the FTV16.5 is acarbon fiber laminate, in this case modeled with 6 layers4. This solution, at equal geometry, results tobe lighter respectively 80% and 60% of the steel and ergal solutions.

Figure 5.15. Initial structural evaluation of different custom main gear struts

Although the laminate solution analyzed is suitable in terms of performance, it results expensive tomanufacture because it requires the design of carbon fibre molds specific for the designed geometry. Twovalid alternatives are sketched in figure 5.16. The concept 1 is made by two central bodies produciblewith 3D printing technique and then bond together through a bath in different layers of carbon fibre.Instead, the concept 2 results from the integration of an off-the-shelf rod in carbon fibre with a 3D printedstrut cover.

Either concepts 1 and 2, can be easily attached to the swivel component, described in section 5.2.1,and to a similar custom swing component on the wheel axle, capable to modify the camber and tow angleof the wheel assembly, since the track dimension is changed at different c.g. locations.

Figure 5.16. Conceptual sketches of the carbon fiber strut

The parasite drag induced by the frontal area of the wheels can be reduced by designing 3D printablefairing covers and wheel pants, aerodynamically shaped.

At last, the landing gear should preserve the integrity of airframe and systems in case of parachuteemergency landing. As already introduced in subsection 5.2.5, the landing gear design is influenced bythe design of the recovery system of the aircraft in order to decide which is the trade-off in terms of rate of

4The layer combination used is [0/ + 45/90/90/ − 45/0]

52


descent to which the landing gear and the recovery system should be designed. The designed emergencysituation corresponds to a vertical sink speed of 15 fps or equivalent impact acceleration around 11.5g. Inorder to not largely oversize the structure and weight of the landing gear system for a parachute landing,the extra kinetic energy, associated with the emergency, is thought to be absorbed by an internal crash-worthy component, located in the middle-low part of the fuselage and connected to the bottom pivotpoints of the two shock absorber.

5.3.2 Nose gear conceptsThe concepts for the nose gear assembly, including the configuration and off-the-shelf possibilities areillustrated in the present section.

Types of nose gear

The first design variable for the nose gear design is the type of configuration. The concepts considered inthis stage are shown in figure 5.17.

Figure 5.17. Sketches of the concepts for the nose landing gear for FTV16.5

The first possibility is a traditional nose landing gear with a single vertical strut mounted directly onthe wheel axle, a setup that guarantees a direct load path to the airframe supports. This solution is thesimplest and cheapest, but the rebound and compression control is exclusively attributed to the shockabsorber, that for this reason needs to have enough stiffness and damping properties to manage all thevibrations and dissipate the related energy.

A valid alternative to the classical straight-vertical nose gear is represented by a trailing link solution.It is made essentially by an L-shaped flexible arm, ahead the wheel, with the shock absorber disposedwith a certain angle with respect to the vertical. In a trailing link connection, the link itself represents anextra shock absorber and contributes to smooth the vibrations caused by hard landing and rough runwaysurface, whilst the shock absorber is responsible to react to the first vibrations induced by landing.An other advantage of the trailing link is the easy maintainability respect to the straight vertical nosegear: repairing the shock absorber from a trailing link gear is quite easy because it does not require todisassembly all the system, but only the shock absorber needs to be replaced. On the other side, a trailinglink means also more moving parts to lubricate and complicates the design of the retraction system, ifrequired, so its choice depends of the type of landing gear for the specific application. [35]

An other important decision, that conditions the design of the nose gear assembly for FTV16.5, is thenumber of tires per strut, choosing from one or two. A double wheel means better ground handling, lowerrolling resistance helping the take-off and redundancy, in case of failure of one assembly, but at the sametime it implies more weight and drag penalties to the aircraft design. Due to the relatively small size ofthe aircraft and low load condition expected on the front of the aircraft, a single wheel unit, with biggerdiameter respect to the scaled down size, can represent a valid alternative capable to have an adequateability to deal with less than perfect runways.

Components off the shelf (COTS)

The initial market research has been done for the wheel and tire applications among big radio commandedairplanes and UAV applications. The feasible options for the concepts described are shown in figure 5.18.

53


Robart tire is a custom goodyear application for small airplane (like scaled Spitfire and P-38) and it has atough outer skin, softer inner core and design for low bouncing. [36] The other solution, used for RC bigjet airplanes (like Skymaster jets), has the same scaled-down outside diameter than the full scale aircraftand so it is a good candidate to be applied, in a double version, for the nose gear of FTV16.5. The tire isrubber type and the sidewalls are slightly thicker than the tread, so it is designed to hold up to side-loadground rolls during steering. [20]

Figure 5.18. Wheel assembly solutions for the nose gear of FTV16.5

The suspension properties are accomplished by a shock absorber to be selected from similar optionsto the ones in figure 5.13, but smaller size and stroke (indicatively 5.5"× 1"). For a straight verticalconfiguration, an oleo shock absorber should be required to mitigate the vibrations induced by bouncingon uneven runways, whilst in a trailing link configuration a simple air-damper spring should be enoughto work in conjunction with the trailing link arm itself.

The nose gear is designed to give to the aircraft the necessary steering ability during ground maneuvers.The steering system on the nose gear can be implemented using a rack and pinion mechanical transmissionor simpler pulleys moved by a belt, as shown schematically in figure 5.19. In both cases, the motionis transfered by a dedicated servo actuator that can be pulsed modulated in order to be commandeddirectly from the autopilot. In the first concept the rotational motion of the nose gear is driven by alinear movement of the rack commanded by an electric motor, while in the second concept the motion istransfered from a smaller pulley with a specific drive ratio. [37]

Figure 5.19. Concepts for the steering system of the landing gear for FTV16.5: Rack and Pinionmechanical transmission on the left and Pulley-Belt mechanism on the right

Custom components

The nose gear is supposed to be positioned at the scaled down longitudinal position respect to the fuselagefront reference and attached to the front bulkhead of the aircraft using a custom bracket in metal. Thestrut and additional scissors to prevents shimmy instability (in case of straight vertical configuration)are addressed to be designed in the preliminary phase of the landing gear design, once the size of theselected components are established.

54


5.3.3 Test rig conceptsThe design of a full landing gear system from scratch requires even the definition of the testing proceduresin order to prove the feasibility of the selected concept and demonstrate the structural integrity uponthe typical take-off and landing conditions, in which the aircraft should operate. Most of the test rigprocedures follows the indications stated in FAR 23 regulations. [12]. The present subsection describesthe concepts and ideas to test the off-the-shelf components, with take-off and landing loading simulatedscenarios, as well as the full assemblies during impact loading.

Shock absorber test rig

Both the selected shock absorbers for the main and nose gear assembly, are needed to be tested in order todetermine the stiffness and damping properties along the full stroke of the piston. Due to the impossibilityto have a shock dynamometer in CfAR shop and to the high cost of using the instrumentation from othershock specialized companies, the first approach to verify the feasibility of the shocks has been to design acustom test rig with a simple leverage mechanism, to be mounted on the precision optical table, providedin CfAR shop. The outlined design sketch and the relative main components are shown in figure 5.20.

Figure 5.20. Test rig for shock absorbers: it shows all the components needed for the test rig assembly.

The system has been designed with a leverage arm ratio of 1 : 8, sufficient to reproduce in the shockabsorber the worst case scenario of impact for design nominal landing. The leverage is realized with adouble rectangular cold steel tube distanced with enough clearance to allow the shock absorber to bepositioned in between and to be moved, when loaded, without interferences. The rectangular cross sectionhas been selected in order to increase the moment of inertia about the longitudinal axis and so decreasethe stress distribution inside the long leverage. The rigidity on the contact area with the shock absorber,pivot point and load input, is increased using specific steel inserts with the same internal shape of therectangular tubes. All the insert in the rectangular bar are fixed with standard screws and hex nuts.

The shock absorber is connected to the two bars with a bolt connection and located in the centerusing two circular steel spacers. The connection of the shock absorber is vertically offset from the topplane on the centerline in order to have the leverage horizontal when the travel of the shock piston is halfstroke: in this way it is valid the approximation of small angles when the bars move. On the other side,respect to the longitudinal centerline, there is house for a counter balance weight used to pre-load andbalance the system, before applying the test loads. The other end of the shock absorber is fixed betweentwo angle brackets designed to be bolted on the optical table provided in CfAR.

The pivot hinge of the leverage system is realized by using a rectangular steel bar in the center planeof the system: it is connected with two pins to additional custom brackets, bolted to the optical table, inorder to remove the oscillations during the motion of the leverage, and different holes in the bottom partare used to create the pivot point of the overall system depending on the height of the shock absorberto test. Even for the pivot steel bar, the central positioning is maintained thanks to round spacers inbetween the two rectangular tubes.

The testing load is applied at the end of the larger side of the leverage using a simple hook, forcetransducer, and a bucket to insert a variable equivalent weight. Two steel spacers are added in betweenthe two bars, in proximity of the hook application in order to increase the bending rigidity of the leveragewhere the load is applied.

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Speed rating test rig

A truck test can be designed to test the speed rating of both main and nose gear selected tires, withequivalent take-off speed of FTV16.5. The test rig can be designed as a trailer for the truck provided inCfAR in order to be easily installable and removable. The basics of the speed rating test rig are shownon the left of figure 5.21. The wheel to test is linked to a fork pivoted on the standard rectangular tubetrailer, used for the connection with the truck. On the same wheel axle, a support plate is pivoted inorder to apply the testing static load on the wheel. The oscillation of the plate is prevented with a doublearm connection with the fork. The equivalent static weight is applied symmetrically on the support platearound the wheel thanks to high density lead rectangular blocks. Safety rubber pads are disposed aroundthe perimeter of the bottom part of the support plate in order to prevent damages on the wheel if therubber tire blows up during the test and to avoid the destruction of the test rig at high speed on theconcrete ground.

If the static test is proved to be successful, the tire can be tested under the real longitudinal spin-uploads, developed on landing when the speed of the tire passes from zero to the aircraft landing speed. Forthis purpose some steel ropes can be used to drop the test rig from a limited height. The ropes shouldbe attached on the wheel axle and commanded from a pulley system inside the truck.

Static drop test rig

A static drop test rig, in compliance with the normative FAR 23 for the height and equivalent mass testparameters, is needed to be designed for testing the main gear and nose gear full assemblies. The testrig for the main gear, sketched on the right of figure 5.21, mainly requires a beefy support for slidingthe double adjustable plate, that represents the support to attach the shock absorbers and to apply thedistributed mass. The bottom plate has two cut-outs in order to allow the vertical motion of the strutsduring impact loading. The double plate before being dropped is connected to the test rig support withhooks and steel ropes with adjustable length according to the required drop height. The sketch in figureshows the overall boundary dimensions needed for a similar test rig concept.

Figure 5.21. Test rig concepts: on the left the Speed Rating Test rig, on the right the Drop Test rig.

5.4 Preliminary DesignIn the preliminary design phase, the landing gear concept is chosen and the design activity becomes moreanalytical and detailed. The designer starts to procure and evaluate the off-the-shelf components available

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in the market and to integrate them into the design. This phase of the project is highly influenced by thedesign of other systems of the aircraft and so it is important to define a design space for the landing gearinside the airframe, in order to avoid all the possible interferences with other components. This phaseof the design includes also the definition of the mechanical subsystems (braking and steering systems),the preliminary design of all the needed structural components and the design of the preliminary testingprocedures in order to determine the feasibility of the system. At the end of the preliminary design stageit is expected that all the landing gear components, with the basic features defined, are placed in CADand integrated with the airframe design.

The present section shows the initial phase of the preliminary design and the planning of the nextsteps needed before entering in the critical design phase of the landing gear, in conjunction with thedevelopment of the airframe design.

5.4.1 Selection of the conceptsThe concepts for the main and nose gear have been selected through an analytical hierarchical approachthat compares the possible solutions described in subsections 5.3.1 and 5.3.2. The most important fea-tures, identified to select the concepts, have been rated with an importance percentage that comes fromengineering criteria and evaluation of the application and integration of the landing gear for FTV16.5.The discriminatory features used for both main gear and nose gear concept evaluation are shown in thecharts in figures 5.22 and 5.23.

Figure 5.22. Feature importance for the Maingear concept selection

Figure 5.23. Feature importance for the Nosegear concept selection

The final selection for the main gear is mostly guided by cost, design simplicity and the needed spaceinside the fuselage, quite limited due to the presence of the center of gravity shifting mechanism andparachute components. The AHP summary table used for the main gear internal concept selection isshown in table 5.7. The selected configuration, with pivot point of the leverage strut on the fuselagebulkhead and the shock absorber directly connected to the strut, has been proved to have the bestfeatures regarding weight, availability and estimated cost, even though the rebound control and stiffnessproperties are all attributed to the shock absorber, unlike the other configurations where the linkagepermits to add an additional kinematic suspension.

Table 5.7. Analytical hierarchical process table for the Main Gear concept selection

Configuration 1 Configuration 2 Configuration 3Features % Rank [1-10] Rank [1-10] Rank [1-10]

Space needed 20 10 7 7Weight 15 8 5 6

Simplicity 20 10 7 5Rebound Control 10 6 8 7

Availability 10 10 8 5Cost 25 8 6 5Tot 100 8.8 6.65 5.85

As for the nose gear concept selection, the final decision is guided by the need to have a gooddirectionality of system and controllable rebound to keep minimum the vibration induced from taxiing

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and take-off run on rough runways. The AHP summary table used for the nose gear concept selection isshown in table 5.8. The trailing link solution, has been preferred to be developed, because it allows touse a simpler air spring shock absorber, since the stiffness and damping are controlled also by the trailingstructure itself.

Table 5.8. Analytical hierarchical process table for the Nose Gear concept selection

Trailing link Classic configurationFeatures % Rank [1-10] Rank [1-10]

Space needed 5 8 6Weight 10 5 8

Directionality 30 9 6Simplicity 5 5 9

Rebound Control 15 8 6Maintainability 10 7 5

Cost 25 9 7Tot 100 8 6.5

The next step, after the concept selection, is to select the off-the-shelf components and integrate theminto the design process. The preliminary CAD models for both main and nose undercarriages are shownin figure 5.24.

The model of the main gear assembly includes the most important components encountered at the cutsection, that corresponds to the main gear positioned in the configuration B (with nominal scaled-downtrack dimension). The basic leverage suspension includes FOX Float DPS with size 6.5” × 1.5”, selectedbecause of its high availability, low cost and the possibility to change the set up of stiffness and dampingbefore each flight. The wheel assembly comprehends the solution 1 for tires, wheels and brakes, describedin subsection 5.3.1, chosen because it fulfills the scaled requirements of FTV16.5 and it’s the simplestand cheapest solution to integrate into the design.

The solid model for the nose gear assembly contains the preliminary brackets that delineate the trailinglink suspension including the same type of shock absorber selected for the main gear, but smaller in size(5.5” × 1”). The wheel assembly solution 1 has been selected, due to the high quality of rubber plusfoam of the tire and because it has a 25% more of the scaled down size, that can help to achieve firstlythe take-off run aircraft speed conditions and absorb the oscillation induced by the runway thanks tothe inner core designed for low bouncing. The steering system represented is a pulley-belt mechanism,selected for its adjustability and relatively ease to be integrated with standard components.

Figure 5.24. Concepts selected for the main gear (on the left) and for the nose gear (on the right)

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5.4.2 Procurement and testing procedures for the off-the-shelf componentsThe present subsection describes the testing procedures necessary to prove the feasibility of the selectedoff-the-shelf components for FTV16.5, even though the tests have not been performed before October 2018due to delays in the procurement of the components and in the manufacturing process of the designedtest rigs.

Wheel assemblies

The combination of tires and wheel selected for both main and nose gear are shown in figures 5.25 and5.26.

Figure 5.25. Procurement of themain gear wheel assembly

Figure 5.26. Procurement of thenose gear wheel assembly

Since the main gear wheel has been designed for not high speed tailwheel applications, it requires speedtesting and certification. Moreover, it doesn’t provide a readily integrated brake system so the wheel canbe considered as starting point for the sizing of the braking system required. At the same time, the nosegear assembly necessitates speed testing, because it has been designed for big radio-commanded airplanesthat usually have lower speed requirements in comparison with FTV16.5.

Both Main gear and Nose gear wheel assemblies are ready to be speed tested, developing the speedrating test rig conceptualized in subsection 5.3.3. The testing procedures needed are summarized asfollows:

1. Weight distribution: the weight distribution on the support plate should represent the maximumvertical loads on the nose and main gear, stated in table 5.4 in subsection 5.2.4. The loadingcondition can be represented by using several layers of four lead blocks, with approximately weightof 10 kg and thickness of 3 mm. The blocks are then locked in position by using steel ropes.

2. Speed-up the truck: the truck should reach gradually the equivalent take-off speed of the aircraft,stated in the list of requirements. If necessary, the tire to test, can be pre-rotated in the oppositesense of the forward speed, by using a simple electric motor, in order to lower the speed requirementon the truck.

3. Evaluation of the tire conditions: after the speed testing evaluate the tread condition of therubber and if visible deformations have been induced from the testing high speed.

4. Dynamic speed test: if item 3 is successful, the next step can be an evaluation of dynamic dropperformances of the tire in order to estimate the spin-up loads induced on the tire when it touchesthe ground at the landing speed of the aircraft. The support plate can be connected, through steelcables, directly to the back of the truck, and all the system can be dropped from a low definedground clearance. The latter can be as minimal as possible in order to reduce the possibility ofblow-up, since no suspension system is used in the speed rating test. The test should only replicatethe phenomena of spin-up, where the tire passes from almost zero speed to the aircraft landingspeed. An indicative value of ground clearance can be 1 inch.

If the selected tires do not pass successfully the described test, bigger aircraft rated and certified solutionsshould be used, as the ones presented in figure 5.10 in subsection 5.3.1.

Shock absorbers

The selected shock absorbers for the suspension system of the nose and main gear are shown in figure5.27. Both solutions have been preferred to an oleo-pneumatic shock absorber because of the less cost,

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weight and space needed inside the fuselage. On the other side, despite of a simple coil spring, theyprovide a manageable compression and damping thanks to the design of two internal chambers where theair passes through, during loading. Figure 5.28 illustrates the cut section view and the most interestingpoints involved in the functionality of the shock absorber: the air is pumped inside the upper chamberby a specific valve (number 4 in figure); the stiffness of the shock can be regulated with a compressionvalve (indicated by number 2 in figure) according to three different setting (firm, normal, soft), that canbe selected depending on the type of runway; the damping is achieved thanks to small orifices like theone shown in item 3; the last item indicates a safety valve of depressurization in order to protect theinternal chambers in case of high load. [38]

Figure 5.27. Off-the-shelf shock absorberfor the nose and main landing gear

Figure 5.28. Internalchambers of the shockabsorber [38]

The shock absorbers are now ready for characterization and for the evaluation and proof of thefeasibility of the selected concepts. The shock load curves have been first determined by an analyticalprocedure that considers the internal geometry of the components shown in figure 5.28.

Although the stiffness and damping effect are correlated in a shock absorber with not separate cham-bers, they have been estimated, via analytical method, as if they act separately. The analytical expressionused to approximate the trend of the spring shock load Fs along the piston travel s, is shown in equation5.13, where Ac is the area of the cylinder, Ap is the area of the piston, V0 is the volume of the upperchamber in static pre-load condition, P0 is the pre-load pressure and γ is the index of the polytropiccompression, assumed to be 1.1 because the reaction is expected to be so fast that it can be assumed tobe close to an isotherm. [39]

Fs = P0Ac

(V0

V0 −Ap · s

)γ(5.13)

On the other side, the damping force has been estimated as a quadratic function of the piston travelrate of change Vt, as shown in equation 5.14, where Ao is the orifice area and Cd is the orifice dischargecoefficient estimated using reference [40].

Fd = ρA3cV

2t

2 (CdAo)2 (5.14)

The resulting shock loads for the shock absorber selected for the main and nose landing gear, infunction of the piston travel, are shown respectively in figures 5.29 and 5.30.

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Figure 5.29. Shock load for the FOX 3 DPS 6.5"×1.5",selected for the Main gear application Figure 5.30. Shock load for the FOX 3 DPS

5.5"×1", selected for the Nose gear application

The true trend of the shock load along the piston travel needs to be validated by experimental tests.Due to the difficulties to use a proper shock dyno test rig that can test the behaviour of the shock absorberin relation to different inputs of force and different setting of the inside pressure, it has been decided totest the shocks in a custom test rig, already described in subsection 5.3.3, with the goal to prove thefeasibility of the shock within the expected impact loads, corresponding to different internal geometries.

The following itemizes summarize all the testing procedures that should be used to characterize theinternal disposition of the shock absorber in FTV16.5.

1. Setting the inside pressure: The inside pressure can be regulated with an hand pump to valuesin between 100 and 300 psi. Note that the shock absorber should be slowly compressed with theshock pump screwed on ten times by around 25% of its stroke, since it is needed equalization ofthe positive and negative air chambers. Typical values of inside pressure are based on the staticload on the shock and are indicated in the manual of the shock absorber.[41] The minimum valueof pressure is set at the beginning of the test in order to balance, with a counterweight, the test rigin the initial position.

2. Setting the SAG: the SAG (negative spring deflection) is the degree by which the shock compressesunder the static load. Typically the SAG can be set in between 20% and 25% of the shock strokefor a firm compression set-up, while a value in between 25% and 35% should be imposed in case ofexpected higher loads, for instance due to the not regular runway. The inside pressure should beregulated until a reasonable value of SAG is achieved. [41]

3. Adjusting the compression: there are three compression settings, with an apposite valve, thatcan be imposed for each test. If the valve is in position "open" the compression is the most sensitiveto unevennesses on the ground, if it is in "drive" it corresponds to the nominal setting of thecompression, whilst in "firm" position the piston is locked and the suspension results more rigid. Ablow-off valve in the bottom part opens the flow of fluid in the case of heavy impacts preventingdamages to the shock.

4. Adjusting the rebound: the rebound can be regulated through a rotating rebound wheel thatchanges the velocity of the piston in loading condition. The rebound can be set depending on thetype of the runway being traveled. If low bouncing is desired the setting wheel valve should berotated in clock-wise direction.

5. Load and stroke data collecting: For each loading testing conditions corresponding to thedifferent inside pressure set, the load on the shock, that results from the leverage ratio and thecorresponding stroke, can be read and collected in an Excel data sheet.

6. Data processing: the data collected can be used to create a curve load/deformation dependingon the variation of the inside pressure.

Note that the testing procedure is applied for different setting of inside pressure and increasing load untilthe SAG requirement is respected.

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5.4.3 Preliminary design of the main gearThe preliminary design of the main landing gear is developed in parallel with the airframe design and eachcomponent should be designed inside a defined volume space, in order not to interfere with the designof other systems inside the fuselage. The design volume is defined considering possible variations of thelength of the internal strut leg, the location of the shock absorber and a certain clearance respect to thesize of the shock absorber. It includes also the central space needed for the design of the crush-worthycomponent for emergency landing. The relative size needed for the preliminary design of the landing gearis shown in figure 5.31. The internal location of the main gear is supposed to be attached on the rearside of the aft bulkhead, since the front side is occupied by the wing attachments and fuel system. Theinternal layout is fixed among the main landing gear 4 configurations and so it needs to be defined forthe worst case landing scenario.

Figure 5.31. Design space for the main landing gear inside the airframe of FTV16.5

The final needed space can be specified after the shock absorber characterization and definition ofthe needed leverage arm ratio between the external (dext) and internal (dint) strut leg. The geometricparameters of the main landing gear in the front plane are shown in figure 5.32. The geometry of the strutis characterized by the determination of the external dext and internal dint lengths, the camber angle γof the wheel, the inclination α of the external strut with respect to a vertical plane, and the inclinationβ of the shock absorber with respect to the internal strut.

Figure 5.32. Preliminary design geometry and load distribution of the main gear selected concept.

One of the results of the shock absorber testing described in subsection 5.4.2 is to find the maximumload that the shock absorber can withstand with the full stroke, at different setting of inside pressure.Once the ultimate load Fs−MAX is determined, it is possible to calculate the required moment arm ratioillustrated in figure 5.32 and so determine the internal layout, when a static load FT−mg is applied to thewheel contact point. The arm ratio can be calculated from an equilibrium momentum about the pivotpoint O that leads to the expression in equation 5.15.

a

b= Fs−MAX

FT−mg(5.15)

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The geometry of the strut should respect the moment arm ratio as described in equation 5.16, whereT is the maximum main wheel track between the possible configurations introduced in subsection 5.2.1and do is the distance of the pivot point from the vertical mid plane of the aircraft, defined at this pointfrom the height requirement of the landing gear and considerations with Airframe Design Team.

a

b=

T2 − do

dint · sin (β) (5.16)

Stated that the values of the track and position of the pivot point cannot be significantly changeddue to the constraint explained in section 3.1, the only chance to validate the use of the selected shockabsorber is to move the length of the internal strut dint and the angle β considering the following designapproach:

1. Change β: The initial value of dint is fixed to the value used to model the initial geometry shownin subsection 5.3.1 and the resulting β is calculated using equation 5.16, checking whether the β isless of 90◦, condition that ensure a valid internal kinematics.

2. Change dint: If the resulting geometry is not valid, fix the angle β to 45◦ and change the length ofthe internal strut, ensuring that the resulting geometry has the internal strut inside the delimiteddesign space. If the geometry is not valid change also the value of the angle β.

3. Change the design space: If items 1 and 2 do not lead to a valid geometry, check if the designspace can be enlarged without hardly interfering the design of other airframe components.

4. Change the shock absorber: If none of the items are respected, the landing gear designer shouldchange the off-the-shelf component and opt for an oleo-pneumatic shock absorber. In that case allthe procedure described should be reiterated for the new component until a valid combination ofshock absorber and internal layout is achieved.

Once the geometry and loads on the shock absorber, in both static and impact cases, are determined itis possible to size the pivot hard-points on the bulkhead using the static equilibrium system of equations5.17. {

Vo = FT−mg + Fs−mg · cos (θ)Ho = Fs−mg · sin (θ)

(5.17)

Preliminary design of the braking system for the main gear

The classical preliminary sizing of the braking system, used to certify braking systems according to FAR23 regulation, starts from the determination of the brake roll distance required to stop the aircraft, asexpressed in equation 5.18. The total landing distance (sb) is determined by the ground roll distancewith brakes applied plus a short "free roll", to account for the reaction time required to apply the brakesand engages the spoilers, that usually is on the order of 1 to 3 seconds. [42]

sb = W/Swρg (CD − µrCL) ln

1 + ρV 2TD

2W/Sw(CD

µr− CL

)+ VTD · tf (5.18)

The landing distance has been calculated for both tailed and tailess configurations of FTV16.5 in orderto consider the worst case scenario for the braking system. The aerodynamic lift and drag coefficientsin landing clean configuration for both tailed and tailess have been scaled from the values of FTV7,according to the data provided by the aircraft aerodynamic team, and are reported, with all the otherneeded values, in table 5.9. The calculations have been done considering no wind and no thrust reversal,with the engine set to IDLE and the friction coefficient with brakes applied equal to µr=0.4. [7]

All the quantities in table are expressed in Imperial units.

Table 5.9. Braking distance calculation according to FAR 23 regulations

W SW VTD z CD CL µR sTOTTailed 386 50.19 83.45 4000 0.0448 0.2894 0.4 585No-Tail 386 50.19 130.90 4000 0.0233 0.0785 0.4 815

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The longitudinal force and the torque needed per each wheel assembly can be calculated assumingthat all the aircraft longitudinal kinetic energy is ideally dissipated by an equivalent longitudinal brakingforce acting along the braking distance, as in the system of equations 5.19.{

12mV

2TD = Fhsb

T = Fh · rw(5.19)

The values for the Kinetic energy and torque needed to size the braking system are calculated andsummarized in table 5.10.

Table 5.10. Calculation of the parameters for the preliminary sizing of the braking system

sTOT Kineticenergy Fh rw T FnTailed 585 42671 ft lbs 71 lbs 1.75 in 124 in lbs 63.6 lbsNo-Tail 815 93000 ft lbs 114 lbs 1.75 in 200 in lbs 102.5 lbs

The table contains even, in the last column, the normal force of the brake pads required to size thebrake disk. In the calculation it has been assumed that the friction coefficient developed on the brakedisk is 0.65 (typical for steel brakes) and the medium disk radius equal to 1.5 inches that guarantees thedevelopment of a brake system with an internal caliper within the wheel size selected, with diameter 3.7inches.

The values required for the braking system sizing have been used to select standard componentsamong the solutions described in subsection 5.3.1. The results show that, if the aircraft is wanted tobe stopped within a specific distance according to FAR regulations, the only chance is to use a brakingsystem certified for ultralight aircraft and in this sense the solution with internal caliper and disk is thepreferable because it is cheap and reliable. One of the possible solutions, with the relative maximumvalues of kinetic energy and torque applied, is shown in figure 5.33. However, to integrate this solutionwith the wheel presented in subsection 5.4.2, it is necessary to design a custom brake disk, since thestandard coupled disk is designed for a wheel larger than the one selected.

Figure 5.33. Braking system for ultralight applications: internal caliper plus brake disk. [43]

In alternative, the complexity and cost of design and production of the braking system could be highlydecreased if it is taken into account that the length of the aircraft is less than 4% of the the total lengthof the runway where it is suppose to fly the aircraft, Foremost (Alberta - Canada). If no requirementson the braking system are applied the aircraft could potentially be stopped in a distance much morelonger than the braking distance sb that results from equation 5.18, and the resulting torque need todissipate the kinetic energy is much more manageable, using hydraulic brakes for rear bike applications.The latter is the strategy often applied to brake UAVs when no specific stopping distances are requiredand the priority is just to brake the aircraft within a defined clearance from the end of the runway, using

64


the simplest, lightest and cheapest solution possible. An example is presented by the solution used forthe experimental aircraft X-48 shown in figure 5.34.

Figure 5.34. Braking system from bike applications applied to the X-48 UAV. [17]

The next step for the design of the braking system is to select an off-the-shelf brake kit from bikeapplications, test it and find the maximum torque applicable, from which it would be possible to estimatethe effective braking distance required and finalize the preliminary design of the braking system.

5.4.4 Preliminary design of the nose gear

A similar approach to the one used for the preliminary design of the main gear has been followed forthe nose gear, starting from the definition of the design space, illustrated in figure 5.35. It has beendetermined from the expected needed space for the steering system and for the attachment to the frontside of the nose bulkhead.

Figure 5.35. Design space for the nose landing gear inside the airframe of FTV16.5

The basic geometry, with the load distribution and the constraints on the pivot point and on the noseattachment point to the front bulkhead, are shown in figure 5.36. As already stated for the preliminarydesign of the main landing gear, the geometry of the nose gear can be finalized once the properties of theshock absorber will be known and tested.

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Figure 5.36. Preliminary design geometry and load distribution for the nose gear selected concept.

The resulting geometric arm ratio, referred to figure 5.36, that defines the final disposition of thetrailing link, can be calculated applying a moment equilibrium about the pivot point of the trailing link,as shown in equation 5.20

c

d= (L1 + L2) · sin (δ)

L1 sin (σ) = Fs−ng

FT−ng(5.20)

where FT−ng is known from the static load distribution between the nose and main gear and Fs−ng is themaximum shock load registered on the shock absorber test. Following an optimization procedure similarto the one described in subsection 5.4.3 for the main gear, the landing gear designer can now rearrangethe inclination of the trailing link respect to the vertical (δ) and the slope of the shock respect to thetrailing link bigger bracket (σ), as well as move the attachment point of the shock on the trailing link(L1 and L2), checking that the geometric requirement 5.20, defined by the load ratio, is respected.

Once the geometry is defined, hard-points corresponding to the pivot on the trailing link and theattachment points on the front bulkhead can be sized using simple static equilibrium equations.

Preliminary design of the steering system for the nose gear

The pulley and belt mechanism has been preferred for the steering system, since it is the most reliable andthe tension applied can be even adjusted with apposite tensioner, in case of harder steering necessitiescaused by rough conditions of the runway. In addition, its maintainability is easier respect to the otherclassical mechanical transmission, since all the units (driven pulley, input pulley and belt) can be replacedsingularly. The initial sizing for the preliminary design of a pulley-belt steering system, sketched in figure5.37, is accomplished defining the center distance of the pulleys and the relative drive ratio. The designapproach, even in this case, has been to consider whether standard sizes of pulleys could satisfy theneeded drive ratio inside a space limited by the volume design, arranged and established in conjunctionwith the airframe design team.

Figure 5.37. Pulley-belt steering system design parameters. [37]

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The calculation of the steering parameters has been done using drive ratio tables contained in reference[37] and the software Center Distance Designer provided by the SDP/SI manufacturer.

Due to the size of the nose gear shaft to steer it has been preferred to have as much contact area aspossible and so the larger pulley, connected to the nose gear and driven by a smaller pulley, whose motionis activated by a dedicated servo. Therefore, the drive ratio results less than one and the transmissionis step-down drive. A value of 0.7 has been chosen as desired drive ratio, since for steering systemapplications on UAVs it is not required an high rate of change of steering. The resulting preliminarysteering system has the parameters shown in table 5.11, where N1, N2 and NB are respectively thenumber of teeth for the larger pulley, smaller pulley and belt.

Table 5.11. Preliminary selection of the pulley-belt steering system

Drive Ratio Center distance Pitch (P) N1 N2 NB0.7 5.0187 in 0.080 30 21 151

The resulting geometry of the steering system is inside the design space and so it results to be avalid candidate for the steering system of FTV16.5. The next step of the preliminary design should bethe procurement and characterization of a servo actuator to connect to the smaller pulley and then theevaluation of forces and torque needed to steer the nose gear shaft within a specific angle ψ with respectto the neutral position.

5.5 Weight, cost estimation and conclusions

The weight of the landing gear is one of the most important factor that, in most of the cases, guides thedesign process and selection of the components. A target of the landing gear weight can be establisheddepending on some basic parameters of the aircraft and of the undercarriage system, using the proceduredescribed in reference [30]. The mass estimation of the system can be calculated according to the empiricalequation 5.21.

WLG = KL ·Kret ·KLG ·WL ·(HLG

b

)· (Ng)0.2 (5.21)

where Kret is equal to 1 for fixed landing gear, KL takes in account the extra landing gear weight neededto land on a carrier aircraft and so in this case is 1, KLG is the landing gear weight factor typically equalto 0.62 for general aviation and home-built aircraft, WL is the landing weight in this case assumed to beequal to the maximum take-off-weight that represents the worst case scenario, HLG is the landing gearheight, b is the wing span, Ng is the selected landing ground load factor. The resulting landing gearweight target is 8 kg that represents 4.5% of the aircraft maximum take-off weight.

The effective weight distribution estimation, at the phase of preliminary design that includes somedefined off-the-shelf components, is shown on the left of figure 5.38. The total estimated weight, based onthe preliminary sizing of basics components and selected standard components, is 9.5 kg. It represents5.4% of the MTOW and is distributed as 79% on the main gear, mostly due to the strut leg designand braking system, and 21% on the nose gear mostly due to the internal steering system and trailinglink structure. As notable the suspension system, for both nose and main gear assembly, results tobe less than 5% of all the landing gear weight, even though the shock absorbers result the principleresponsible components to dissipate the kinetic energy developed on landing. This first estimation ofweight distribution can tend better towards to the target with the optimization of the design, successiveto the detailed definition of all the components needed.

The cost estimation of the landing gear has been conducted through a market analysis of the typicalcosts of components needed and purchased. [31] The most huge amount of cost is expected to be reservedfor the production of the main gear strut legs, especially if they require custom mold design. The secondsystem that affects the cost budget is the suspension system. Braking system and steering system havesimilar cost estimation and the effective value depends upon the use of standard components.

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Figure 5.38. Weight and Cost estimation for the production of one landing gear set for FTV16.5

Conclusions

The undercarriage system, at this point, results defined in all its basic functionality and the detailedsizing of all the needed components can be started as soon as the off-the-shelf components are testedaccording to the described testing procedures. The process requires close collaboration with the airframedesign team in order to delineate the hard-points inside fuselage bulkhead. From the first definition of thelanding gear layout it is important to have an idea of the performance of the aircraft with undercarriage onthe most critical phases for a landing gear point of view: take-off and landing. A preliminary evaluationof the basic performances on take-off and landing of a simplified model of landing gear, will be presentedfor both 7% and 16.5%, in the next chapter.

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Chapter 6

Basic performance evaluation ofaircrafts with a landing gear system

The behaviour of the aircraft on ground in most of the cases is not linear, mainly due to the complexityof the landing gear in order to absorb energy on landing and dampen vibrations during the groundmaneuverability. Dynamic loads are developed and their determination is important to evaluate theperformance of the landing gear and aircraft itself during all the phases of preparation and flight. Thebasic performance evaluation is treated in the present chapter, assuming the gear as a linear or non-linear black-box depending on the type of loads developed on the particular phase analyzed. The basicdynamic response is evaluated for both FTV7 and FTV16.5 equipped with the undercarriages designedand described in chapters 3 and 5.

The first dynamic phase analyzed is the taxiing/take-off run over a typical runway surface model,used to certify, according to FAR regulations, the dynamic response of the aircraft with landing gear: thenecessity to evaluate the response of the aircraft with landing gear over an uneven runway is highlightedby the results of ground mock-ups for FTV7, as described in chapter 4. The other dynamic responseevaluated for aircrafts with landing gear is the landing, assumed to be a two-point touch-down, as usuallyperformed from aircraft with conventional tricycle landing gear.

6.1 Response of the aircraft with landing gear during the typicalground maneuvers

The taxiing response considered in this thesis work, is concerned to the entire phase of ground movementstowards to the straight line motion on ground before the final take-off, without including braking andturning that are considered as separate operations. The dynamic calculations and simplifications usedfollow the rational criteria that are common used for the design of landing gear for bigger aircraft, inorder to meet the certification requirements. [?]

6.1.1 Taxiing dynamic modelDuring taxiing the non linear response of the aircraft in terms of heave and pitch motion, is mostlydue to the non linearity of the runway above which the landing gear is demanded to move the aircraft.In order to empathize the irregularities of the ground for the dynamic response, the landing gear hasbeen modeled as a simple linear spring/damper system. The reaction on the undercarriages has beenevaluated considering the aircraft as a rigid body, the worst case scenario in terms of induced vibrationsand bouncing on ground.

Since every runway has its own profile, in a design phase, the aircraft response with landing gear isevaluated over a typical non-smooth runway, the San Francisco 28R1, according to the CS-25.491. [?]The runway is modeled as a series of different dips and bumps with 1-cosine shape, as shown in figure6.1.

1San Francisco 28R is often used to evaluate the performance of the landing gear still in a design phase, so often itis called "Design runway".

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Figure 6.1. Aircraft taxiing operation over a 1-cosine modeled runway

The modeling of the runway depends upon the depth h (xr) of dips/bumps at the location xr respectto the reference coordinate system, usually located where the irregularity starts. The detailed procedureto model the runway is described in Appendix C. The dynamic model of the aircraft response is evaluatedby the two degrees of freedom most involved in this phase, heave and pitch motion, applying the Lagrangemethod, as discussed in the Appendix C. The system of second order differential equations, shown in 6.1,has been reduced to first order differential equations and solved using the time-space domain in Matlab®& Simulink. The degree of freedoms are evaluated at the center of gravity position and the equivalentmatrices that describe the system are dependent on the position of the nose and main gear with respectto the center of gravity and equivalent properties of linear stiffness and damping for the main and nosegear. [

M]{z

θ

}+[C]{z

θ

}+[K]{z

θ

}=[CR]{hN

hM

}+[KR

]{hNhM

}(6.1)

Note that the non linearity on the response is imposed by the right-hand side terms in the system ofdynamic equations due to the variation in elevation of the runway surface.

6.1.2 Taxiing performance evaluation and comparison for FTV7 and FTV16.5The time-history response of the aircraft during taxiing over a typically bumpy runway has been evaluatedfor both FTV7 and FTV16.5 and the major results, in terms of vertical acceleration and heave response,are illustrated and commented in the present subsection. The mass, geometry, stiffness and dampingproperties of the FTV7 and FTV16.5, used to solve the dynamic model are explicated in table 6.1.

Table 6.1. Input for the taxiing dynamic model

m Iy lN lM KN KM CN CM[kg] kgm2 [m] [m] [N/m] [N/m] [N s/m] [N s/m]

FTV7 13.6 2 0.474 0.0526 6500 32067 50.726 229.25FTV16.5 175 75 1.117 0.124 27450 82138 555 2508

The mass (m) and location of nose (lN ) and main gear (lM ) for both FTVs results from the scalingprocess of the full scale aircraft, with the center of gravity located at 56% of the main aerodynamicchord. The pitch inertia for FTV7 has been estimated using the approximate inertia process described insubsection 4.1.1, whilst for FTV16.5 the value has been obtained from the inertia load cases stated in theflying demo spec document in reference [2]. The linearized stiffness for the main and nose landing gearof FTV7 are the ones obtained from ground testing described respectively in subsections 3.2.4 and 3.3.5.For FTV16.5 instead, the stiffness properties result from the linearization of the polytropic theoreticalrelation used for a preliminary characterization of the shock absorbers, as described in subsection 5.4.2.The damping properties for FTV16.5 have been estimated using a similar linearization, and the resultingvalues have been scaled for FTV7 using the MTOW and stall speed formal factors.

The two aircrafts have been evaluated when they encounter a 1-cosine bump, with depth and lengthrespectively equal to ∆ (h) = 15mm and LR = 8m for FTV7, and ∆ (h) = 30mm and length LR = 15mmfor FTV16.5, when the aircrafts are at 75% of the take-off speed. The values for the geometric modeling

70

6 – Basic performance evaluation of aircrafts with a landing gear system

of the runway, have been estimated from the real condition of the airstrip in Merrit - Douglas Lake forFTV7, while for FTV16.5 typical values used for ultralight aircraft have been found in reference [39].

The heave acceleration and displacement response at the nose and main landing gear’s location, forthe two flight test demonstrators, are respectively shown in the graphs a and b of figures 6.2 and 6.3.

Figure 6.2. Vertical acceleration during taxiing over a 1-cos dip runway for the FTV7 and FTV16.5

Figure 6.3. Heave response during taxiing over a 1-cos dip runway for the FTV7 and FTV16.5

The most relevant observations for the taxiing performance evaluation when the aircraft encountersa bump/dip, are underlined in the following itemize:

• Rebound induced by the bump: the negative acceleration and displacement values in the figuresare caused by the rebound of the nose and main landing gear that occurs when they encounter theground, that can be considered an infinite rigid target.

• Maximum value of heave acceleration: for FTV7 the maximum value (0.55g) is registered onthe nose gear when it encounters the bump, while for FTV16.5 the maximum (0.63g) is observableon the first rebound of the nose gear, probably due to the nose gear shock absorber recoil duringrebound. As observable the maximum acceleration transmitted to the landing gear structure andthen transferred to the airframe, is almost the same for the two FTVs, since the two landing gearshave been designed with structural similarities.

• Maximum value of heave displacement: the heave response follows the motion induced by thebouncing acceleration caused by the encounter of a bump inside the runway. The maximum peakvalue (18mm for FTV7 and 40mm for FTV16.5) is determined by the depth of the bump and thecompression when the rigid target is encountered.

• Peaks and oscillations: the peaks are a measure of the rebound control performed by the landinggear. The peaks of the main gear follow the ones of the nose gear with a lag that is dependenton the location of the main gear with respect to the nose gear. For FTV16.5 the heave motion is

71


completely damped in 5 peaks within 1.5 seconds after the bump encounter, while the motion isnot properly damped for the FTV7 since no specific damping components are used. Note that thebouncing effect on FTV7, could be significantly reduced by adopting tires with larger diameter andso higher surface to damp vibrations.

In last measure, the analysis of the results is voted to understand the applicability of the adopted modelsand the reliability of the assumptions made. The linearization assumptions are fully true only for thenose gear of FTV7 since it is characterized by a linear progressive spring. The real behaviour of themain landing gear for FTV7 and both undercarriages of FTV16.5 is not linear since the suspensionsystem is respectively made by a leaf strut, with behaviour similar to a parabola, and air/damper shockabsorber with exponential behavior weighted by the index of gas compression inside the shock, similarto an isotherm reaction. On the other side the damping properties of the FTV16.5 shock absorber aredependent on the square of rate of position change of the piston, whilst for FTV7 the only damping isprovided by the rubbers on the main gear and by inertial effect. Anyway, the simplified model explainedin this section highlights the basic taxiing response of the aircraft empathizing the importance of the nonlinear geometry of the runway for the performance evaluation.

6.2 Response of the aircraft with landing gear during a typicaldesign landing

Landing is the most critical phase for both landing gear and airframe design, because a significantamount of energy needs to be dissipate in order to reconduce safely the aircraft on ground. The landingphase analyzed in this section is the operation typically performed with the aircraft flaring and the nosegear still airborne at the moment of landing. The rates of descent used for both FTV7 and FTV16.5performance evaluation are the same used for the design, that come from previous flights for FTV7 andFAR regulations for FTV16.5.

6.2.1 2-points landing dynamic modelA simple representation of the aircraft during touch-down on the main landing gear, is to consider thewheel/tire assembly and the suspension system in series, distinguishing the aircraft mass m from theunsprung mass mT (that consists of wheels, tires, brake assemblies on the two wheels). The multi-system model analyzed is sketched in figure 6.4. The dynamic behavior of the landing gear is describedby the equivalent stiffness KM and damping cM of the suspension system and an equivalent stiffnessKT of the tire/wheel assembly. In this analysis, the main gear strut has been considered rigid, thesuspension system aligned with the aircraft center of gravity and the tire without damping properties.These assumptions have been done since, for a nominal landing of FTV 16.5, it is assumed that all theimpact load is transferred to the internal shock absorbers, the center of gravity is quite close to theground contact point on the main landing gear and the damping forces on the tires are expected to besmall in comparison with the damping properties of the shock. [39]

Figure 6.4. Aircraft model during landing

The aircraft displacement and tire compression are described by the quantities zM and zT , whilethe compression on the suspension system is expressed as zSA = zM − zT . The forces acting on theaircraft at the moment of the touch-down are the weight, residual lift and the non linear load on theshock absorber induced by the ground reaction inside the landing gear system, and for simplicity in thismodel they have been assumed to be applied on the aircraft center of gravity. The landing gear, on

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6 – Basic performance evaluation of aircrafts with a landing gear system

a landing dynamic analysis, needs to be considered non linear due to the non linear nature of impact.The dynamic behaviour is governed by the system of equation 6.2, where the last equation represent theinitial condition of vertical speed of the mass and tire, assumed to be equal to the design vertical rate ofdescent VTD at the instant before the touch-down.

mzM + fNL (zSA, zSA) +mg − L = 0mzT − fNL (zSA, zSA) +KT zT = 0zM (0) = zT (0) = VTD

(6.2)

The system of dynamic equation are modeled and solved in Simulink, whose block diagram is shown inAppendix C.

6.2.2 Landing performance evaluation and comparison for FTV7 and FTV16.5The landing performance evaluation is measured in terms of impact acceleration developed and consequentwheel travel and shock absorber piston’s motion. The non linear properties of the landing gear have beenintroduced in the model thanks to look-up tables: the stiffness and damping properties for FTV16.5are taken into account thanks to shock load graphs introduced in subsection 5.4.2; as for FTV7, theexperimental non linear load-displacement curve, described in subsection 3.2.4, is imported in order torepresent the structural response of the leaf strut. The stiffness effect on the tire has been assumedto be linear and estimated considering the maximum rolling radius upon MTOW loading distribution.The other necessary inputs (weight properties and geometry) are the same used for the taxiing model,introduced in table 6.1.

The impact acceleration and vertical displacement time-history responses are shown in figures 6.5 and6.6

Figure 6.5. Vertical impact deceleration on the main gear of FTV7 and FTV16.5

Figure 6.6. Vertical displacements on landing on the main landing gear of FTV7 and FTV16.5

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The most valuable results deduced from the graphs are itemized as follows:

• Tire effect: Note that adding a simple tire model for both FTV7 and FTV16.5 the accelerationdoes not tend to infinite values at the moment of impact.

• Peak of impact acceleration: the impact acceleration is a function of landing weight and of thevertical rate of descent before the aircraft touches down. The maximum peak value for both FTVs iscomparable with the design load factor assumed to be 3 for both landing gear design cases. In termsof structure response it is important the values above the design load factor, if present, are limitedto a small amount of time so that they cannot induce permanent and catastrophic deformations:as notable on the graphs, this is respected for both FTV7 and FTV16.5.

• Vertical bouncing: The vertical bouncing on the main gear in this analysis does not consider theeffect of the nose gear that rotates and touches the ground.

• Displacements: the wheel travel (25mm) consequent to impact landing is quite similar to the oneobtained in the ground drop test (27.4mm) shown in subsection 3.3.5. In the case of FTV16.5, thedisplacement consequent to impact landing, is distinguished between the wheel travel and piston’smotion of the shock absorber. The large wheel travel (the maximum value is 130 mm) is due tothe nature of the suspension system finalized to have the maximum gear displacement controlledby the shock absorber’s stroke. The shock absorber in the model, has been considered in order tohave the full stroke when the impact acceleration presents a peak.

Although the dynamic landing model is a good mean to verify the aircraft response with landing gear ina phase of preliminary design, it needs to be improved including the real geometry of the landing gear,that in the case of FTV16.5 can be finalized only when the internal location of the shock absorbers isdetermine. For the best accurate response of the aircraft with landing gear, separated main gear modelsand nose gear effect should be considered in the analysis, and the model should be integrated with thecomplete Simulink model of the aircraft: this approach can lead to the performance evaluation of all the9 degrees of freedom involved on the aircraft during landing.

74

Chapter 7

Conclusions and Futuredevelopments

7.1 ConclusionsThe design and development from scratch of a fully operative landing gear, applied in this dissertation fortwo different scales of aircraft, lead to several causes for reflection that highlight the complexity related toexperimental design. The nature of Unmanned Aerial Vehicles and their applications require iterations,ground testing and eventual redesign of components in order to improve and optimize the final layoutand behaviour of the landing gear.

Both FTV7% and FTV16.5% applications, evidence that the landing gear design for flight test de-monstrators of bigger aircrafts, goes far beyond merely scaling down the size of the full scale solution.Although the designer has not wide freedom to change some parameters that define the final geometryand disposition of the undercarriages, he needs to have in mind that the resulting behaviour on impactlanding and the complexity of the suspension system depends on the specific scale of the aircraft and onthe operational conditions of the runway. The small scale of FTV7% does not allow a geometric scaling ofwheel assemblies and so an additional requirement of maximum 25 % of extra size, as stated in subsection3.1.4, has been established with the stakeholders. Furthermore the need to test the feasibility of take-offand landing operations, within a large range of center of gravity, has not allowed the positioning of themain landing gear at the scaled down value, because the take-off requirement, stated in section 2.1, and areasonable load ratio distribution between the nose and main landing gear, should be respected for eachc.g. position.

The design for both aircrafts has been conditioned by cost, weight, availability and manufacturabilityconstraints. As introduced in the current state of the art of landing gear for similar aircraft applications insection 2.3, the market research for aircraft off-the-shelf components that alleviate the cost of a customsolution, has been significantly complicated passing from the small scale FTV7% to FTV16.5% andsolutions from other field of applications have represented the best alternative, even though appropriatetests are required to define the feasibility and integration with the overall design.

The specific design of the landing gear for the small scale FTV has underlined the difficulties tointegrate a new system into an already built airframe, especially because the available space inside thebays of the aircraft has not been designed to house the internal components of the nose steering systemand attachment plates of the main gear leaf strut and so design trade-off and redistribution of internalcomponents have been constantly necessary. Once the landing gear has been installed into the airframesof the two tailed and no-tail configurations of FTV7%, new ground mock-ups have been needed in orderto validate the behaviour of the aircraft with undercarriages on the runway, and improve the design ofsome components (tire size and horns for the steering system), before proceeding straight to a flight test.

The design of the undercarriages for FTV16.5% has shown up that the landing gear is not a "Cinderellasubject" but, in some cases, it becomes a significant design driver, influencing the external and internallayout of the airframe since the initial phases of conceptual and preliminary design of a new aircraft.Therefore close collaboration with the other design teams of the aircraft (airframe, recovery system,system controls) has been maintained along all the crucial design decisions, avoiding possible interferencesand inconsistencies.

The two designed landing gears have been finally evaluated in terms of heave response during themost critical phases of take-off over a bumpy runway and landing. The irregularities on the two runways

75


chosen as model for the ground application of FTV7 and FTV16.5, as discussed in chapter 6, producesimilar average heave acceleration in the two landing gear structures, even if the induced oscillations resultalmost no damped for the 7% and well damped for the 16.5%, due to the different suspension systemimplemented. As for the heave response transmitted after touching the ground in a design landing, thetwo landing gears receive a similar normalized acceleration because both of them are designed to sustainnominally a 3G landing, but the resulting motion on the two applications, although they have a similartrend, is characterized by a different travel distribution depending on the weight of the aircraft and onthe suspension system used.

7.2 Future developmentsThe FTV7%, in both tailed and no-tail configurations, is awaiting for its first flight test campaignwith landing gear. The flight tests with variable position of the main landing gear and the successivedata analysis process are already planned, as discussed in subsection 4.3.1. The flight test results canvalidate the structural response of the landing gear, estimated in the initial phase of the design, usingstructural simulation tools, and then verified by ground tests. Furthermore they can underline whatis the performance of the aircraft with landing gear during take-off and landing at different position ofcenter of gravity: the response evaluation can lead to the validation of the developed model describedin chapter 6 and to the definition of the best position of c.g. with the preferred control setting for theautopilot, that can support further design decisions for the landing gear of FTV16.5%.

As for the landing gear of FTV16.5% the progress in the design will follow the development of theairframe of the aircraft, expected to be ready for production in 2019. The preliminary design can continuefollowing the procedures described in subsections 5.4.3 and 5.4.4, once the shock absorbers for both mainand nose gear will be characterized after testing, according to the steps defined in subsection 5.4.2. Thepreliminary phase of the project will be ultimated once all the aspect introduced in chapter 5 will befinalized, including the definition of all the components and design parameters of the needed mechanicalsubsystems (steering and braking system) and the definition of the required test rigs and procedures toverify the behaviour of the landing gear assemblies.

In the next stages the design will be refined and finalized in every detail so that all the parts willbe geometrically and structurally fully determined and ready to be manufactured. In this phase, all thepossible failure modes during parachute emergency landing and the resulting risk management will beconsidered, in conjunction with the Recovery System Team.

If no redesign will be required, the component list definition and assembly CAD drawings will beprepared and the project will move to the procurement and readiness phase with the landing gear ma-nufactured and assembled. After that the assembled landing gear will be ready to be ground tested andintegrated with the first developments of the airframe. The whole aircraft, with landing gear, is thenscheduled to be ready to fly in summer 2019.

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Appendix A

Landing gear Matlab parametercalculator

The MatLab® code, attached to the present appendix, computes the basic parameters and features of thelanding gear depending to some input that can be defined by the user at the beginning of the program.It is an useful tool that can be used for the initial sizing of the landing gear, where some parameters arenot fully defined and iterations are necessary.

A.1 Input parametersThe input parameters, that the landing gear designer has to define according to a list of requirementslist, are:

• Aircraft and performance data:

– Maximum Take of Weight MTOW;– Mean aerodynamic chord M.A.C. length and location;– Wing reference area;– Rate of descent;– Aircraft configuration: tailed or no tail;

• Geometry input:

– Location of the nose gear;– Center of gravity position;– Wheel track for the main gear;– Size of the wheel assemblies;

• Design choices:

– Ground reaction load factor Ng1;– Vertical load ratio between nose gear and main gear;

A.2 Output parametersThe output data computed by the program are:

• Geometry output:

– Location of the main gear;

1It is the ratio between dynamic and static loads, used to estimate the ground reaction loads during impact

77


– Geometry angle:∗ Angle A between the C.G. and the vertical line on the main gear;∗ Angle C between the C.G. and the vertical line on the nose gear;∗ Tip-back angle αtb;∗ Angle αM−N between the nose gear and main gear outboard wheel;∗ Turnover angle ψ;

• Impact parameters:

– Wheel travel of the landing gear;– Normalized acceleration developed on landing, depending on the rate of descent and compres-

sion of the strut;– Longitudinal spin-up and spring-back loads developed on landing;

• Other design parameters:

– Take-off Rotation requirement verification;– Braking system sizing:

∗ Brake kinetic energy;∗ Torque required per each wheel;∗ Horizontal and vertical load during braking;

– Shock absorber sizing:∗ Stiffness load;∗ Damping load;

– Drop Test parameter;

The code, applied for FTV16.5, is listed as follows.

1 clc2 clear all345 %%FIXED INPUT67 MTOW =175.24; %% [kg]8 S_ref =50.19; %%[ft ^2]9 MAC =1.5697 e+03; %% [mm] Mean aerodynamic chord from Horizon -16.5% BA

document10 x_MAC =886.46; %% [mm] Leading edge of the MAC11 x_nose =647.70; %% [mm] Location from BBA - scaled down dimension12 track_main =1388.71; %% Scaled outboard wheel distance for the Main Gear

calculated from BA CAD13 V_land =39.93; %% [m/s]14 g =9.81;1516 %% VARIABLE INPUT1718 %%1 PARAMETER : CG POSITION19 percentage_cg_MAC =58;2021 %%2 PARAMETER : LOAD RATIO22 percentage_load_main =90;23 percentage_load_nose =10;24 load_ratio = percentage_load_nose / percentage_load_main ; %% Nose gear Load

over Main Gear Load2526 %%3 PARAMETER : WEIGHT

78

A – Landing gear Matlab parameter calculator

27 W=MTOW;2829 %%4 PARAMETER : TAIL OR NO TAIL30 tail =1; %% 1= tail configuration , 0=no tail configuration3132 %%5 COMPRESSION33 d_scaled =64.11/1000; %% Scaled down compression [m]; The actual

compression is 92.29 mm34 d_design =92.29/1000;35 d =0.0300:0.0001:0.100;3637 %%6 RATE OF DESCENT38 V_z_FAR =7.33*0.3048; %% [m/s] from FAR 23. 15 fps from Parachute landing

.39 V_z_parachute =15*0.3048;4041 %%7 Design Ground load Factor42 N_g =3; %% Ground Reaction landing factor to estimate vertical dynamic

loads , suggested by FAR 23.434445 %% GEOMETRY CALCULATOR4647 alpha=asin (( V_z_FAR / V_land ))*180/ pi;48 V_horizontal = V_land *cos(alpha*pi /180);4950 x_cg =( percentage_cg_MAC /100* MAC)+x_MAC; %% [mm] Longitudinal position of

the c.g.51 z_nose =346.55; %% [mm] Vertical location of the nose reference to the

ground reference line52 z_cg= z_nose +56.90; %%[mm] Vertical loacation of the c.g. with respect to

the ground reference line. Max value , from document BOMBARDIER SPECREV D

5354 l_n=x_cg - x_nose ; %% [mm]. Position of the nose gear with respect to C.G.55 l_m=l_n *( load_ratio ); %% [mm]. Position of the main gear with respect to

C.G.56 x_main =l_m+x_cg; %%[mm] Longitudinal position of the m.g.5758 wheelbase =l_n+l_m;5960 A=( atan(l_m/z_cg))*180/ pi; %% Angle between the vertical line on the

main gear and the line to the C.G.61 C=( atan(l_n/z_cg))*180/ pi; %% Angle between the vertical line on the

nose gear and the line to the C.G.62 if(tail ==1)63 x_tip =3496.84; %% [mm] Location of the tip and farther point of

the aircraft . Data from Solidworks model64 x_tip_contact =x_tip - x_main ; %% [mm] distance between the contact

point and tip of the tail65 y_tip_contact =433.45; %% [mm] vertical distance between the

contact point and the tip of the tail66 else67 x_tip =3038.87; %% [mm] Location of the tip and farther point of

the aircraft . Data from Solidworks model68 x_tip_contact =x_tip - x_main ; %% [mm] distance between the contact

point and tip of the tail

79


69 y_tip_contact =373.39; %% [mm] vertical distance between thecontact point and the tip of the tail

70 end7172 alpha_tipback =( atan( y_tip_contact / x_tip_contact ))*180/ pi; %% Tipback

angle , calculated using dimensions form the Solidworks model73 alpha_main_nose =( asin (( track_main )/(2*( wheelbase ))))*180/ pi; %% In plane

angle between the Nose gear and the outboard wheel74 alpha_turnover =atan(z_cg /( l_n*sin( alpha_main_nose *pi /180)))*180/ pi; %%

Turnover angle in the cross plane. the maximum Turnover angleaccording to FAR 23 is 55 degree

7576 %% LOAD CALCULATOR7778 % Static Loads79 P_m =(W*l_n)/(2*( wheelbase )); %% [kg] Static Load on each wheel of the

Main Gear80 P_n=W*l_m /( wheelbase ); %% [kg] Static Load on the Nose Gear81 P_m_max =N_g*P_m; %% [kg] Peak vertical load on each wheel at landing82 P_n_max =N_g*P_n; %% [kg] Peak vertical load on the nose gear at landing8384 static_loads = [P_m P_n ];85 maximum_loads = [ P_m_max P_n_max ];8687 % Impact Loads88 a_g_FAR_design = V_z_FAR ^2/(2* d_design *g); % average acceleration developed

during landing , considering a compression of 92.29 mm89 a_g_parachute_design = V_z_parachute ^2/(2* d_design *g); % average

acceleration developed during parachute landing vonsidering acompression of 92.29 mm

9091 % Fixed Rate of descent92 a_g_parachute_scaled = V_z_parachute ^2./(2* d*g);93 a_g_FAR_scaled = V_z_FAR ^2./(2* d*g);94 %95 plot(d*1000 , a_g_parachute_scaled )96 hold on97 grid MINOR98 plot(d*1000 , a_g_FAR_scaled )99 xlabel ('Compression [mm]');100 ylabel ('Normalized acceleration [a/g]');101 title('Normalized acceleration depending on the wheel travel ');102 legend ('Rate of descent : 15 fps [ Parachute Landing ]','Rate of descent :

7.33 fps [FAR Calculation ]');103104 % Fixed Compression105 d_scaled =64.11; % [mm]106 d_CAD =92.29; %[mm]107 d_FAR =84.8; %[mm]108109 V_z =0:0.1:15; % [fps ]. 15 fps is the parachute landing110 a_g_scaled =( V_z .*0.3048) .^2./(2*( d_scaled /1000) *g);111 a_g_CAD =( V_z .*0.3048) .^2./(2*( d_CAD /1000) *g);112 a_g_FAR =( V_z .*0.3048) .^2./(2*( d_FAR /1000) *g);113114 figure115 plot(V_z , a_g_scaled )116 hold on

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117 plot(V_z , a_g_FAR )118 plot(V_z , a_g_CAD )119 grid MINOR120 xlabel ('Rate of descent [fps]');121 ylabel ('Normalized acceleration [a/g]');122 title('Normalized acceleration depending on the Rate of Descent ');123 legend ('Scaled down compression : 64.11 mm ', 'Nominal Landing Compression

: 84.8 mm', 'Parachute Landing Compression : 92.29 mm');124125 %% TAKE -OFF ROTATION REQUIREMENT126 % Altitudes127 T_0 =288; %% Temperature at z128 h=6.5;129 z =0.887; %% Foremost airportaltitude in [km]130 z_c=z+0.3; %% cruise altitude at 200 m above the RWY131 rho_0 =1.225; %% Air density at sea level [kg/m^3]132 rho=rho_0 *(( T_0 -h*z)/T_0) ^4.2561; %% Air density at the airport height133 rho_c =1.0146; %% Air density at cruise height134135 % Speeds136137 V_S =32.46; %% [m/s] Stall speed at at MTOW clean configuration [From

aerodynamic team]138 V_R =1.2* V_S; %% Rotation speed at TO: 1.1 to 1.3 V_S139 V_c =52; %% [m/s] Cruise speed from spec document BBA]140141 % Geometric data wing and tail142 AR=b^2/ S_ref; %% Aspect Ratio143 S_h =0.67075; %% Tail surface : from BBA documentation : 7.22 sqft144 x_acwf =(0.25* MAC) +34.9*0.0254; %% aerodynamic center wing fuselage : 25%

MAC , Leading edge at 34.9" of the F.S. [ according to the documentHorizon -16.5 _Flying_Demo_Specs_RevB ]

145 x_ach =3.1056; %% aerodynamic center tail:from BBA documentation : 122.27in

146 z_T =0.637; %% Arm of the Thrust with respect to the ground referenceline [From Solidworks 16.5%]: center of the Nacelle

147 z_nose =0.412; % %177.22 is the height of the Nose station with respectthe RWY

148 z_cg =0.46914; %% vertical position of the c.g. Maximum position from thedocument [0066 - BBA 16.5% LANDING GEAR SYSTEM ].

149 z_D=z_cg; %% Arm of the Drag with respect to the ground reference line.It has been assumed that Z_D equal z c.g.

150 x_cg =0.56* MAC +34.9*0.0254; %% forward limit of the c.g. at 56% MAC151152 % Aerodnamic coefficients153 C_D0TO =0.0668; %% C_D0 at TO with flaps and U-Tail154 e =0.92; %% Ostwald factor155 K=1/(e*pi*AR);156 C_LC =(2*W)/( rho_c*V_c ^2* S_ref); %% C_L cruise157 C_Lflap =1; %% extra_lift coefficient at TO due to flap. No Flap are for

the GEN2B and GEN2C158 C_LTO=C_LC+ C_Lflap ; %% C_L at TO159 C_D= C_D0TO +K*C_LTO ^2; %% C_D at TO160 C_macwf = -0.005; %% C_m0 wing -body161 C_Lh = -1.1; %% Negative Lift coefficient Tail/ elevon at TO162163 % Forces164 D=0.5* rho*V_R ^2* C_D*S_ref; %% Total Drag at TO

81


165 M_acwf =0.5* rho*V_R ^2* C_macwf *S_ref*MAC; %% Pitching moment wing -body atthe a.c.

166 L_TO =0.5* rho*V_R ^2* S_ref*C_LTO; %% Total Lift at TO167 L_h =0.5* rho*V_R ^2* S_h*C_Lh; %% Lift Tail168 L_wf=L_TO -L_h; %% Lift wing fuselage at TO169 N=W-L_TO; %% Normal force on the Main gear at TO170 F_F=mu*N; %% Friction force171 a_x =(T-D-F_F)*g/W; %% longitudinal acceleration [m/s2]172173 % Calculation of the longitudinal position of the main gear with respect

to the FWD limit of C.G. [ Reference Mohammad , Sadraey - AircraftDesign & Systems Engineering Approach ]

174 I_yy =90.718; %% Inertia of the aircraft w.r.t. C.G175 I_yymg =I_yy; %% Inertia of the aircraft at TO with respect to the ground

contact point at the main gear176 theta_ddot =15* pi /180; %% required rotation acceleration at TO [deg/s^2]:

For a normal light general aviation aircraft it is 8-10 deg/s^2177 x_mg_loop =( I_yymg *theta_ddot -D*z_D+T*z_T+M_acwf -(W/g)*a_x *( z_cg)-W*x_cg+

L_wf*x_acwf -L_h*x_ach)/( L_wf+L_h -W);178 x_mg =1.905;179 tail =1;180 if(tail ==1)181 while(abs(x_mg - x_mg_loop ) >1e -6)182 I_yymg =I_yy+W/g*(x_mg -x_cg)^2;183 x_mg_loop =x_mg;184 x_mg =( I_yymg *theta_ddot -D*z_D+T*z_T+M_acwf -(W/g)*a_x *( z_cg)-W*

x_cg+L_wf*x_acwf -L_h*x_ach)/( L_wf+L_h -W);185 end186 else187 L_h =0;188 while(abs(x_mg - x_mg_loop ) >1e -6)189 I_yymg =I_yy+W/g*(x_mg -x_cg)^2;190 x_mg_loop =x_mg;191 x_mg =( I_yymg *theta_ddot -D*z_D+T*z_T+M_acwf -(W/g)*a_x *( z_cg)-W*

x_cg+L_wf*x_acwf -L_h*x_ach)/( L_wf+L_h -W);192 end193 end194195 %% SHOCK ABSORBER SIZING196197 %Data198 mh = m / 2; % Factor total mass for

half aircraft199 Wh = mh * 9.81;200 Lh = Wh; % Lift balances weight201 W_e = 2.23; % Vertical speed on

landing (TAS)202203 % Single main landing gear parameters leading to non - linear behaviour204 Pinf = 3.4 e5; % Extended position pressure [50 psi]205 PS = 5.4 e5; % Static position Pressure206 PC = 2.4 e6; % Compressed position pressure207 PA = 1e5; % Atmospheric pressure208 zS = 0.0381; % Maximum Stroke [m]209 Area = 1.6e-3 %Wh / PS;210 Vratio = PC / Pinf;211 Vinf = Vratio * Area * zS / ( Vratio - 1);212 zinf = Vratio / ( Vratio - 1) * zS;

82


213 ns = 1;214 nd = 1.35;215216 dz = 0.001; z = [0: dz: zS]; [dummy , nz] = size(z);217 for j=1: nz218 Pd(j) = Pinf / (1 - z(j) / zinf)^nd;219 Fd(j) = (Pd(j) - PA) * Area;220 end221222 plot(z,Fd);223 hold on224 grid on225 xlabel ('Displacement ');226 ylabel ('Force ');227228229 %% BRAKING SYSTEM SIZING230 distance_landing =250; %%[m]. Use Mohammed231 t_stop =(2* distance_landing )/ V_horizontal ;232 a_braking = V_horizontal / t_stop ;233 F_stop =MTOW* a_braking ;234 KE =0.5*( MTOW)* V_horizontal ^2; % KJoule235 KE_req_per_wheel =(KE /(0.737562) )/2; % Ft lbs236237 tire_main_diameter =6; %% inches238 Torque =(( MTOW /2)* a_braking )*(( tire_main_diameter *25.4/1000) /2); %% [N m]

Torque required per each wheel239 Torque_req = Torque *8.851; %% [in lbs] Torque required per each wheel240241 F_N_braked = static_loads (2)+W* a_braking *( z_cg)/( wheelbase ); %% Vertical

loads on the Nose gear with brakes applied242 F_M_braked =(( static_loads (1))*2-W* a_braking *( z_cg)/( wheelbase ))/2; %%

Vertical loads on the Main gear with brakes applied243 braked_loads =[ F_N_braked F_M_braked ];244245 %% DROP TEST PARAMETERS246 W_e=MTOW *0.454;247 h =(3.6*( W_e/S_ref)^0.5); % FAR 23 [mm]

83

84

Appendix B

CAD assembly drawings

B.1 Assembly drawings and Bill of components for the LandingGear for FTV7%

This appendix shows all the assembly drawings that define the bill of components used for the landinggear for FTV7. The main gear drawings are illustrated in figures B.1 and B.2, while the nose assemblydrawing are illustrated in figures 4.2, B.3 and B.4

B.2 Assembly drawings for the Shock Absorber Test RigThe shock absorber Test Rig described in subsection 5.3.3 of chapter 5, is ready for manufacturing andits CAD assembly drawing with all the components needed is shown in figures B.5 and B.6.

85


X.XX5X.XXX ANGLES:

MACHINING REQUIREMENTS:

±0.02"

±0.005"±0.002"

+0.005"-0.002"

±0.005"BETWEEN HOLES:

- BREAK SHARP CORNERS AND EDGES 0.005" MAX.- INTERNAL CORNERS OF CUTOUT R0.010" MAX.

HOLE ≤1":

GENERAL TOLERANCES:

DIMENSIONS ARE IN INCHESUNLESS OTHERWISE SPECIFIED

X.X

DO NOT SCALE DRAWING

±0.1°- CHAMFER EXTERNAL THREAD 45 TO MINOR DIA. - C'SINK INTERNAL THREAD 90 TO MAJOR DIA.

X.XX ±0.005"

R63N7

1

2

3 4

5

5

6

7

3

Item No. Part Number Description Quantity1 0060-0206 Main gear strut 1

2 97231A440 Locknut for 10-32 screw 6

3 91525A115 Washer for 10 screw size 12

4 0060-0209 Support attachment plate 2

5 1296N21 Rubber sheet 46 Aircraft Skin Aircraft Skin 17 92200A347 Socket head Screw

10-32, 1" long 6

This part is designed to be made by

CHECKED

DRAWN

NAME DATE

Fabrizio Rizzi

[Kg]; [Lbs]

111:2

DWG NO.

MATERIAL: DIMENSIONS ARE IN

PARENT ASSEMBLY

WEIGHT:

FINISH:

TITLE:

BSIZE

SCALE: ofSHEET:

DWG REV

APPROVED

THE INFORMATION CONTAINED IN THISDRAWING IS THE SOLE PROPERTY OFUVIC CENTRE FOR AEROSPACE RESEARCH (CfAR). ANY REPRODUCTION IN PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF UVIC CfAR IS PROHIBITED.

PROPRIETARY AND CONFIDENTIAL

CfAR OFFICE NUMBER778-351-1926

R:\0066-BOMBARDIER-16.5% LANDING GEAR\DESIGN\MECHANICAL\7 %\7%_CAD_LANDING_GEAR\SWEPT_VERSION\FULL

FILE REV

Component List exploded view 4A A

B B

C C

D D

E E

F F

8

8

7

7

6

6

5

5

4

4

3

3 2 1

Figure B.1. Main Gear assembly drawing 1: leaf strut and attachment with the airframe

X.XX5X.XXX ANGLES:


±0.02"

±0.005"±0.002"

+0.005"-0.002"



HOLE ≤1":

GENERAL TOLERANCES:


X.X



X.XX ±0.005"

R63N7

18

910

10

11

12

13

Item No. Part Number Description Quantity1 0060-0206 Main Gear Strut 18 Tire and wheel

70mmTire and wheel

70mm 1

9 92981A042 Shoulder screw 5mm, M4x0.7mm 1

10 91116A140 Washer for M5 screw 2

11 92871A042 Spacer for M5 screw, 6mm long 1

12 92871A031 Spacer for M5 screw, 2 mm long 1

13 94710A101 Locknut for M4x0.7 mm 1 This part is designed to be

made by CHECKED

DRAWN

NAME DATE

Fabrizio Rizzi

[Kg]; [Lbs]

111:1

DWG NO.


PARENT ASSEMBLY

WEIGHT:

FINISH:

TITLE:

BSIZE

SCALE: ofSHEET:

DWG REV

APPROVED




R:\0066-BOMBARDIER-16.5% LANDING GEAR\DESIGN\MECHANICAL\7 %\7%_CAD_LANDING_GEAR\SWEPT_VERSION\MAIN_GEAR_ASSEMBLY_WITH_SWEPT

FILE REV


B B

C C

D D

E E

F F

8

8

7

7

6

6

5

5

4

4

3

3 2 1

Figure B.2. Main Gear assembly drawing 2: wheel assembly connection with the leaf strut

86

B – CAD assembly drawings

X.XX5X.XXX ANGLES:


±0.02"

±0.005"±0.002"

+0.005"-0.002"



HOLE ≤1":

GENERAL TOLERANCES:


X.X



X.XX ±0.005"

R63N7

1

2

35

67

3

8

4 Item No. Part Number Description Note

1 93985A209 Shoulder screw 3/16", 1-1/4"

2 90945A760 Washer 1/4"3 0060-0220 Bushing for the Rim 3 is connected to

the rim

4 Nose Gear Piston strut

5 92311A321 Set screw 4-48

6 0060-0221 Bushing for the Nose gear strut

6 is connected to the nose gear strut

7 Nose gear Tire and Rim

8 90633A009 Locknut 8-32


CHECKED

DRAWN

NAME DATE

Fabrizio Rizzi

[Kg]; [Lbs]

111:2

DWG NO.


PARENT ASSEMBLY

WEIGHT:

FINISH:

TITLE:

BSIZE

SCALE: ofSHEET:

DWG REV

APPROVED





FILE REV


B B

C C

D D

E E

F F

8

8

7

7

6

6

5

5

4

4

3

3 2 1

Figure B.3. Nose gear assembly drawing 2: wheel assembly and piston strut

X.XX5X.XXX ANGLES:


±0.02"

±0.005"±0.002"

+0.005"-0.002"



HOLE ≤1":

GENERAL TOLERANCES:


X.X



X.XX ±0.005"

R63N7

24

15

28 29

30

31

32

33

34

35

25

37

Item No. Part Number Description Quantity15 0060-0210 Support Plate for

the Nose Gear 1

24 2444 Linkage for the steering system 1

28 0060-0225 3D printed support for the servo 1

29 HV6130 Servo for the steering system 1

30 20170919001 Coupler for the servo 1

31 91221A112/95868A132

Glass-Filled Nylon screw/ nylon screw

2-561

32 0060-0219 Bushing for the steering shaft 1

33 0060-0217 Steering shaft 134 0060-0212 Horn for the

steering shaft 1

35 92196A109Socket Head

Screw 4-40, 7/16" long

1

25 98019A216 Washer for N4 137 91834A102 Nut for 4-40 screw 1


CHECKED

DRAWN

NAME DATE

Fabrizio Rizzi

[Kg]; [Lbs]

111:1

DWG NO.


PARENT ASSEMBLY

WEIGHT:

FINISH:

TITLE:

BSIZE

SCALE: ofSHEET:

DWG REV

APPROVED





FILE REV


B B

C C

D D

E E

F F

8

8

7

7

6

6

5

5

4

4

3

3 2 1

Figure B.4. Nose gear assembly drawing 3: steering system

87


X.XX5X.XXX ANGLES:


±0.02"

±0.005"±0.002"

+0.005"-0.002"



HOLE ≤1":

GENERAL TOLERANCES:


X.X



X.XX ±0.005"

R63N7

1

23

6

4

5

7

8

12

11

10

9

Item No. PN Description QTY1 Shock 6.5" 12 A0001-0001 Rect. Bar 23 A0001-0003 Insert 2 24 A0001-0004 Vertical Pivot Plate 15 A0001-0005 Angle for the bar 26 A0001-0006 Angle for the shock 27 1/2-20 screw 28 92196A389 1/2-20 screw 19 90128A961 #10-32 screw 110 92390A908 Pin 3/8" 111 98306A809 Pin 3/8" 212 91864A062 1/4-20 screw 8


CHECKED

DRAWN

NAME DATE

A0001_1

[Kg]; [Lbs]

111:20

DWG NO.


PARENT ASSEMBLY

WEIGHT:

FINISH:

TITLE:

BSIZE

SCALE: ofSHEET:

DWG REV

APPROVED




R:\0066-BOMBARDIER-16.5% LANDING GEAR\DESIGN\MECHANICAL\16.5 %\TEST RIGS\SHOCK_ABS_TEST\A0001

FILE REV

DR. AFZAL S.Exploded view shock Test rigA A

B B

C C

D D

E E

F F

8

8

7

7

6

6

5

5

4

4

3

3 2 1

Figure B.5. Shock absorber Test Rig assembly drawing 1

X.XX5X.XXX ANGLES:


±0.02"

±0.005"±0.002"

+0.005"-0.002"



HOLE ≤1":

GENERAL TOLERANCES:


X.X



X.XX ±0.005"

R63N7

2

13

9

7

Item No. PN Description QTY2 A0001-0001 Rect Bar 27 1/2-20 screw 19 90128A961 #10-32 screw 113 A0001-0002 Insert 2 2


CHECKED

DRAWN

NAME DATE

A0001_2

[Kg]; [Lbs]

11

DWG NO.


PARENT ASSEMBLY

WEIGHT:

FINISH:

TITLE:

BSIZE

SCALE: ofSHEET:

DWG REV

APPROVED




R:\0066-BOMBARDIER-16.5% LANDING GEAR\DESIGN\MECHANICAL\16.5 %\TEST RIGS\SHOCK_ABS_TEST\A0001

FILE REV

DR. AFZAL S.Exploded view shock test rigA A

B B

C C

D D

E E

F F

8

8

7

7

6

6

5

5

4

4

3

3 2 1

Figure B.6. Shock absorber Test Rig assembly drawing 2

88

Appendix C

Performance evaluation

The dynamic response of the landing gear is almost controlled by two dynamic components in series, thetire and shock absorber. The overall system should be properly modeled in order to have an accuratedistribution of loads when the aircraft is in operation. The behaviour of tires under static and dynamicvertical and later loads is complex and depends upon material properties, pressure and temperature, wallflexibility and runway conditions. [44] A simple representation of the tire can be an undamped modelwith stiffness proportional to tire pressure.

C.1 Taxiing Model

The landing gear, in the first phases of preliminary evaluation of concepts and off-the-shelf componentscan be assumed to be a simple linear spring/damper, resulting in a set of linear equations. The taxiingresponse analyzed, refers to irregularities on the runway, that can be modeled as a series of dips andbumps, in a 1-cos shape.

Each dip/bump can be modeled according to the equations C.1, where ∆hr and Lr are respectivelythe depth and length of dip or bump, xr is the distance along the runway, h (xr) and h (xr) are theelevation profile and rate of change at xr coordinate, assuming a true air speed V at the instant of timeconsidered. [39]

h (xr) = ∆hr2

(1 − cos 2πxr

Lr

)(C.1a)

h (xr) = dh

dxr

dxrdt

= Vdh

dxr(C.1b)

The nose and main gear pass at different times over the dip/bump causing different heave and pitchresponse of the aircraft. The motion of the aircraft, considered as a rigid body, over a one dip-cosinerunway described in equation C.1 is regulated by the response of the landing gear, simply modeledwith two mass-spring-damper1 respectively for the nose and main gear. The overall model, analyticallyanalyzed and implemented in Matlab®/Simulink refers to the figure C.1.

1The spring and damping coefficients are the ones resulting from a single model respectively for the nose and maingear.

89


Figure C.1. Rigid aircraft with equivalent model of undercarriages during taxiing over the design runway.[39]

The two degree of freedom that represents the response of the aircraft at the center of mass are theheave zc (downwards positive) and pitch θ (relative to any horizontal datum). The aircraft responses areevaluated with respect to a datum initial state, so they are incremental. The data that play an importantrole in the aircraft response are:

• Mass properties of the aircraft:

– Mass of the aircraft: m– Inertia about the center of mass: Iy

• Stiffness properties of the landing gear:

– Equivalent stiffness on the nose gear: KN

– Equivalent stiffness on the main gear: KM

• Damping properties of the landing gear:

– Equivalent damping on the nose gear: cN– Equivalent damping on the main gear: cM

• Geometry of the landing gear:

– Distance of the nose gear from the c.g.: lN– Distance of the main gear from the c.g.: lM

The dynamic equation that describe the motion can be obtained from the Lagrange”s equations, shownin C.2, with the generalized coordinates zc and θ.

d

dt

(∂T

∂xj

)− ∂T

∂xj+ ∂D

∂xj+ ∂U

∂xj= ∂ (δW )∂ (δxj)

for j=1,2 (C.2)

The kinetic energy, potential energy and dissipation function for the generalized model considered areexpressed in equation C.3.

T = 12mzc

2 + 12Iy θ

2 (C.3a)

U = 12KN∆N

2 + 12KM∆M

2 (C.3b)

D = 12CN ∆N

2 + 12CM

˙∆M2 (C.3c)

90

C – Performance evaluation

The motion equations of the 2DoF system, shown in C.4, are obtained substituting the equation C.3 intothe equation C.2. [

m 00 Iy

]{Zcθ

}+[

CN + CM −lNCN + lMCM−lNCN + lMCM lN

2CN + lM2CM

]{˙zCθ

}(C.4)

+[

KN +KM −lNKN + lMKM

−lNKN + lMKM lN2KN + lM

2KM

]{zCθ

}={f(t)lT f(t)

}The system of equations can be applied for a dynamic system like the one described in chapter 6 inequation 6.1 and solved with time-space method in Matlab® & Simulink. with time-space method

C.2 Landing simulink model

Figure C.2. Simulink model for the heave response on time domain during landing

91

92

Bibliography

[1] C. H. Wolowicz, J. S. Bowman-Jr, and W. P. Gilbert, “Similitude requirements and scaling relati-onships as applied to model testing,” NASA Technical Paper, 1979.

[2] Bombardier and Q. Aerospace, “Horizon - Flying Demo Specs for the new generation FTV,” 2018.

[3] CfAR and Q. Aerospace, “BBA 7pcnt (Operational Documentation and Scaling process),” 2018.

[4] F. Rizzi, CfAR, Bombardier, and Q. Aerospace, “Landing Gear project charter for the 16.5% FTV,”2018.

[5] H. G. Conway, “Landing Gear Design,” 1958.

[6] A. Argiolas, “Landing Gear Design and integration into aircraft,” 2016.

[7] J. Roskam, Airplane Design. 1985.

[8] N. S. Currey, Aircraft Landing Gear Design: Principles and Practices. 1988.

[9] M. C.-Y. Niu, Airframe Structural Design - Practical Design Information and Data on AircraftStructure. 1988.

[10] Mohammad H Sadray, “Aircraft Design - A systems engineering approach,” 2013.

[11] V. N. Divakaren, G. V. V. Ravikumar, and S. R. Patnala, “Aircraft Landing Gear Design andDevelopment,” 2017.

[12] Federal Aviation Authority, “Chapter 13 - Aircraft Landing Gear Systems,” Aviation MaintenanceTechnician Handbook - Airframe, 2014.

[13] Al-Hussaini, “Undercarriage Layout Design,” 2015.

[14] A. Jha, “Landing Gear Layout Design for Unmanned Aerial Vehicle,” 14th National Conference onMachines and Mechanisms (NaCoMM09), NIT, Durgapur, India, December 17-18, 2009, 2009.

[15] G. Roloff, “Aircraft Landing Gear: The Evolution of a system,” in Airbus-Deutschland GmbH,no. April, 2002.

[16] T.-U. Kim, J. Shin, S. Kim, and I. Hwang, “Design of a crashworthy landing gear using compositetube,” ICCM International Conferences on Composite Materials, 2009.

[17] B. Management, “X-48B Blended Wing Body Flight Control Demonstrator,” 2009.

[18] Federal Aviation Administration, “Far Part 23 — Airworthiness Standards,” 2010.

[19] U. Commander John R. Brown, A critical study of Spin-up drag loads on aircraft landing gears. PhDthesis, 1949.

[20] Morne and Duke, “Skymaster jets ARF PLUS - Technical Manual,” 2015.

[21] Platform and U. Subsystems, “PENGUIN C - Unmanned Aircraft System,” 2017.

[22] R. F. Woodbury, Elements of Parametric Design. 2010.

[23] Hobbyking, “Aircraft RC parts catalog US,” 2018.

93


[24] J. Moriarity and C. Gallagher, “The Ultimate Guide to Waterjet,” Flow Internal Corporation, 2001.

[25] S. Benson, “Bending Basics: The fundamentals of heavy bending,” The Fabricator, 2017.

[26] A. Handbook, “Heat treating of Aluminum Alloys,” ASM International, vol. 155, 1999.

[27] Pyrotek Aerospace, “Certificate of conformance for Heat treatment of 7075-T6 Aluminum,” 2018.

[28] S. T. Chai and W. H. Mason, “Chapter 3 Landing Gear Concept Selection,” in Landing GearIntegration in Aircraft Conceptual Design, 1997.

[29] M. P. R. Products, “Airplane series Servo catalogue,” 2017.

[30] Krauss and Wille, “Chapter 8 - Weight Estimation,” 2002.

[31] CfAR and Quaternion, “Documentation of Purchase orders relative to the landing gear componentsfor FTV7%,” 2018.

[32] B. CfAR, “2018-08-08 FTV2B TAIL BFP Test Info,” 2018.

[33] F. Rizzi, “BBA 16.5%FTV core requirements for the landing gear,” 2018.

[34] Marc-Ingegno, “Analisi strutturale FEM – Cerchio scomponibile per carrello di velivolo,” iMex A,2017.

[35] R. Mark, “Basics of Trailing-Link landing gear,” 2017.

[36] Robart, “Catalogue and specifications for tires and wheels for small airplane applications,” 2018.

[37] SDP/SI, “Master inch catalog D820: drive components used in mechanical transmission,” 2018.

[38] MTBR, “The Fox Float RP2(3) damper service thread: uncovering the secret,” 2017.

[39] J. E. C. Wright, Han R., Introduction to aircraft Aeroelasticity and loads. 2014.

[40] I. Neihouse, W. Pepoon, L. Aeronautical, and L. A. Force, “National advisory committee for aero-nautics,” 1950.

[41] DT Swiss AG, “Dt Swiss - SHOCKS Technical Manual,” 2015.

[42] Warren F. Phillips, Mechanics of Flight. Wiley, 2009.

[43] MATCOmfg, Wheels and Brakes W600, W600XT, W600XLT. 2016.

[44] H. B. Pacejka, Tyre and Vehicle Dynamics. 2006.

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