Scuola di Ingegneria
Dipartimento di Ingegneria civile e industriale
More-Electric Aircraft Landing Gear Actuation
System: Models for Preliminary Design
Tesi di Laurea Magistrale in Ingegneria Aerospaziale
Candidato:
Marco Ramagini
Relatori:
Prof. Ing. Eugenio Denti
Prof. Ing. Gianpietro Di Rito
Tutor:
Ing. Ulrich Kling
Bauhaus Luftfahrt e. V.
Università di Pisa
Dipartimento di Ingegneria Civile e Industriale
Bauhaus Luftfahrt e. V. Taufkirchen, Germany
More-Electric Aircraft Landing Gear Actuation
System: Models for Preliminary Design
Master Thesis in Aeronautical Engineering
Candidate: Marco Ramagini
Academic Supervisors: Prof. Ing. Eugenio Denti Prof. Ing. Gianpietro Di Rito
External Supervisor: Ing. Ulrich Kling
Ringraziamenti
Arrivato al termine di questo lungo percorso, è giusto guardarsi indietro e ringraziare tutte le persone che lo hanno reso possibile.
Vorrei ringraziare il mio relatore al Bauhaus Luftfahrt, l’Ing. Ulrich Kling, per la grande disponibilità che mi ha dimostrato per tutta la durata di questa tesi. Senza di lui questo lavoro non avrebbe preso vita.
Un ringraziamento importante va ai miei relatori qui a Pisa, i Proff. E.Denti e G.Di Rito. Vorrei ringraziarli non solo per il sostegno dimostratomi durante la stesura di questa tesi ma anche per la conoscenza trasmessami in questi anni.
Sono infinitamente grato ai miei genitori per aver reso possibile la scrittura di queste righe. Ringrazio mia Mamma, per tutti i sacrifici che ha fatto e per aver sempre creduto in me. Mi ha reso quello che sono diventato oggi. Ringrazio mio Babbo, per le parole di incoraggiamento che hanno preceduto ogni esame e per avermi insegnato a non arrendermi mai.
Un ringraziamento speciale va a mio fratello, Juri, la persona più importante della mia vita. Questi anni di convivenza, fatti di risate e litigi, resteranno indimenticabili. Ovunque io sarò potrai sempre contare su di me.
Voglio ringraziare i miei amici di una vita, Berte e Profè. Sebbene viviamo in tre città diverse e ognuno abbia preso la propria strada, siamo sempre rimasti uniti. Mi avete dato la forza per arrivare in fondo. Ringrazio in particolare Berte, per esserci stato nei momenti più difficili: sei l’amico migliore che si possa desiderare.
Vorrei ringraziare Leo, anche se non riusciamo a vederci spesso, so che posso sempre contare su di te, nonostante i 10000 km che ci separano.
Ci tengo a ringraziare tutte le persone che hanno reso questi anni di Università indimen-ticabili: Malco, Paolo, Leo, Ja, Cami, Lisa, Andre, Giampy e Gioele. Un ringraziamento speciale lo devo a Umbè e Fili, con i quali tutto è iniziato. Infine voglio ringraziare il Manfro, sono onorato di aver affrontato questo percorso insieme a te.
Abstract
The growing interest in environmental issues has led the industrial policies to turn their attention on the research of more eco-friendly solutions. More-Electric Aircraft concept of using electricity as the only power source for non-propulsion purposes is an effective mean for achieving this goal. This explains why in recent years a great effort has been made in order to develop Power-by-Wire actuators. Even if the conventional centralized hydraulic system has proven to be a solid solution for decades, it offers limited potential for improvements and uses non-environmentally friendly fluids to perform actuation. Thus, electric actuators are becoming increasingly attractive for removing the natural drawbacks of conventional actuation.
In a context where the flight control actuation is moving towards more electric solu-tions, there is no longer reason to employ the conventional hydraulic system for landing gear actuation, which is the largest user of hydraulic power in civil transport aircraft. This thesis aims to provide the designer with models able to asses landing gear actuation by means of an Electro-Mechanical Actuator (EMA) and an Electro-Hydrostatic Actu-ator (EHA). In this way the designer will be able to make top-level decisions between competing solutions based on a greater and deeper knowledge of the system.
To achieve this goal, first, information about the state of the art of EMA and EHA technology has been gathered and the main requirements regarding landing gear actuation have been derived. Secondly, the models of a landing gear, an EMA and an EHA were developed using Modelica language which is especially suitable for simulating complex multi-physical systems. At the same time, an automatized user-OpenModelica interface has been implemented using Java language, in order to make the models re-usable for a rapid exploring and assessment of different architectures.
Simulation results have highlighted the superiority of EMAs in terms of power con-sumption. However, EMA need a more sophisticated control system mainly to decelerate the actuator at the stroke end. An actuator mass estimation pointed out that the EMA is an interesting solution for a overall weight reduction.
Finally, the interface developed in Java has proven to be a suitable mean for the implementation of optimization algorithms for the two actuators design.
Sommario
Il crescente interesse per le tematiche ambientali ha portato le politiche industriali a rivol-gere la loro attenzione alla ricerca di soluzioni più ecocompatibili. Il concetto di More Electric Aircraft di utilizzare l’energia elettrica come unica fonte di energia per scopi non propulsivi è un mezzo efficace per raggiungere questo obiettivo. Questo spiega perché negli ultimi anni è stato fatto uno sforzo notevole per sviluppare attuatori Power-by-Wire. Infatti, anche se il sistema idraulico centralizzato convenzionale ha dimostrato di essere una soluzione solida e affidabile per decenni, offre un potenziale limitato di miglioramento e utilizza fluidi non rispettosi dell’ambiente per l’azionamento. Di conseguenza, gli attua-tori elettrici stanno diventando sempre più attraenti per la rimozione degli inconvenienti naturali dell’azionamento convenzionale.
In un contesto in cui l’azionamento dei comandi di volo si sta orientando verso soluzioni più elettriche, non c’è più ragione di utilizzare il sistema idraulico convenzionale per l’azionamento del carrello di atterraggio, che è il maggiore utilizzatore di potenza idraulica nei velivoli da trasporto civile. Questa tesi ha lo scopo di mettere a disposizione del progettista dei modelli in grado di valutare le prestazioni di un azionamento dei carrelli di atterraggio tramite un attuatore elettro-meccanico (EMA) e un attuatore elettro-idraulico (EHA). In questo modo il progettista sarà in condizione di prendere decisioni di alto livello tra soluzioni concorrenti basate su una maggiore e più profonda conoscenza del sistema.
Per raggiungere questo obiettivo, in primo luogo, sono state raccolte informazioni sullo stato dell’arte della tecnologia EMA ed EHA e sono stati ricavati i principali requisiti relativi all’azionamento dei carrelli. In seguito, i modelli di un carrello di atterraggio, un EMA e un EHA sono stati sviluppati utilizzando il linguaggio Modelica, che si presta particolarmente alla simulazione di sistemi multifisici complessi. Parallelamente, è stata implementata un’interfaccia utente-OpenModelica utilizzando il linguaggio Java, al fine di rendere i modelli riutilizzabili per una rapida esplorazione e valutazione di diverse architetture.
I risultati delle simulazioni hanno evidenziato la superiorità dell’EMA in termini di assorbimento di potenza. Tuttavia, l’EMA necessita di un sistema di controllo più sofisti-cato che permetta di decelerare l’attuatore in corrispondenza del fondo corsa. Una stima della massa degli attuatori ha evidenziato che l’EMA è una soluzione interessante per una riduzione complessiva del peso.
Infine, l’interfaccia sviluppata in Java ha dimostrato di essere un mezzo per la imple-mentazione di algoritmi di ottimizzazione per la progettazione dei due attuatori.
Contents
1 Introduction 1
1.1 Motivation . . . 1
1.2 Objectives . . . 3
1.3 Approach and Methods . . . 3
1.4 Outline . . . 4
2 Literature Review and Requirements 5 2.1 State of the Art . . . 5
2.1.1 Introduction . . . 5
2.1.2 Research Programmes on EHA . . . 5
2.1.3 Research Programmes on EMA . . . 9
2.1.4 Electric Motors . . . 11
2.1.5 EMA Mechanical Transmission . . . 12
2.1.6 EHA Hydraulic Circuit . . . 15
2.2 Design Requirements . . . 15
2.2.1 Introduction . . . 15
2.2.2 Load Requirements . . . 15
2.2.3 Time Actuation Requirements . . . 17
3 Landing Gear Actuation System 19 3.1 Landing Gear Modelling . . . 19
3.1.1 Introduction . . . 19
3.1.2 OpenModelica Models . . . 22
3.2 Electro-Mechanical Actuator Model . . . 24
3.2.1 Introduction . . . 24
3.2.2 Electric Motor . . . 25
3.2.3 Mechanical Transmission . . . 40
3.2.4 Sensor Models . . . 43
3.2.5 Control System . . . 44
3.3 Electro-Hydrostatic Actuator Model . . . 46
3.3.1 Introduction . . . 46
3.3.2 Electric Motor . . . 48
3.3.3 Hydraulic Circuit . . . 48
3.3.4 Control System . . . 54
3.4 Control System Design . . . 55
3.4.1 Electro-Mechanical Actuator . . . 55
3.4.2 Electro-Hydrostatic Actuator . . . 57
3.5 Lock System Model . . . 58
CONTENTS CONTENTS
3.7 Complete Models . . . 60
4 Interface in Java Environment 63 4.1 Introduction . . . 63
4.2 Modify Model XML file . . . 63
4.3 Simulation Run . . . 65
4.4 Simulation Results . . . 66
4.5 Optimization Algorithm . . . 66
4.6 Mass Estimation . . . 68
4.6.1 EHA Mass Estimation . . . 68
4.6.2 EMA Mass Estimation . . . 70
4.6.3 Method for Mass Estimation in Java . . . 71
5 Practical Case Study 73 5.1 Introduction . . . 73
5.2 Parameters Setup . . . 73
5.2.1 Landing Gear . . . 73
5.2.2 Electro-Mechanical Actuator . . . 75
5.2.3 Electro-Hydrostatic Actuator . . . 77
5.2.4 Actuators Control System . . . 77
5.2.5 Lock System . . . 81
5.2.6 Aerodynamic Force . . . 81
5.3 Simulation Parameters Setting . . . 81
5.4 Simulation Results . . . 82
6 Conclusion and Outlook 93 6.1 Conclusions . . . 93
6.2 Future Work . . . 94
A Landing Gear Data 103
List of Figures
1.1 Schematic representation of the interface implemented in Java Environment. 3 2.1 Actuator evolution from centralized hydraulic circuit to Elecetro-Mechanical
actuator [1]. . . 6
2.2 IAP working scheme [1]. . . 7
2.3 Landing Gear Steering Actuator [1]. . . 8
2.4 EHA architecture and overall view of the EHA-FD of THERMAE II pro-gram [1]. . . 8
2.5 EMA demonstrator [1]. . . 11
2.6 Comparison between roller and ball screw contact area [2]. . . 13
2.7 Different Motor-Screw layouts [1]. . . 14
2.8 Landing gear retraction time [3]. . . 17
3.1 Landing gear main components. . . 20
3.2 Landing gear components [4]. . . 20
3.3 Example of retraction mechanisms [5]. . . 21
3.4 Force-stroke graph for efficiency calculation [5]. . . 22
3.5 Stick-diagram of the retraction mechanism. . . 22
3.6 OpenModelica model of the first retraction mechanism. . . 23
3.7 Implementation of the second retraction mechanism. . . 24
3.8 EMA operating principle scheme [6]. . . 25
3.9 EMA implementation in OpenModelica. . . 25
3.10 Schematic representation of a brush motor (left) and a brushless motor (right) [7]. . . 26
3.11 Working scheme of a BLDC motor. . . 26
3.12 Example of winding configurations [8]. . . 28
3.13 OpenModelica model of the BLDC motor. . . 29
3.14 Schematic representation of a two phase motor [8] . . . 30
3.15 One phase on operation in two a two phase motor [8]. . . 31
3.16 Schematic representation of a two phase on operation motor [8]. . . 32
3.17 Schematic representation of an H-bridge device [8]. . . 32
3.18 Current decay in an H-bridge circuit. . . 33
3.19 Back-emf force in a three phase motor [8] . . . 34
3.20 Three phase on operation scheme [8]. . . 34
3.21 Working scheme of a three phase motor and its six step driver [8]. . . 35
3.22 OpenModelica model of the 3-phases brushless DC motor and drive. . . 35
3.23 Representation of the hysterisis PWM technique [8]. . . 36
3.24 Representation of the Clocked Turn-On PWM technique [8]. . . 37
LIST OF FIGURES LIST OF FIGURES
3.26 PWM technique. . . 38
3.27 Motor friction torque as function of motor speed [9]. . . 39
3.28 Comparison between the two electric motor models. . . 40
3.29 Schematic representation of the three parts the screw model can be divided in [10]. . . 40
3.30 Example of friction models [11]. . . 41
3.31 Friction-displacement relation as according to Rabinowicz [11] . . . 42
3.32 Simplified friction model [12]. . . 42
3.33 EMA typical control scheme [13]. . . 44
3.34 EMA closed loop system. . . 45
3.35 EMA Control System. . . 45
3.36 Inverse model block. . . 46
3.37 EHA working principle [6]. . . 47
3.38 Hydraulic circuit in OpenModelica. . . 48
3.39 OpenModelica pump model. . . 49
3.40 Ideal single ended cylinder [14]. . . 50
3.41 OpenModelica cylinder model. . . 51
3.42 Working principle of a mechanical-hydraulic cushion [15]. . . 51
3.43 Cylinder leakage flow. . . 53
3.44 EHA control system model. . . 55
3.45 Actuator force as a function of α . . . 56
3.46 Pump variables. . . 58
3.47 Cylinder variables. . . 58
3.48 Lock System model in OpenModelica. . . 59
3.49 Aerodynamic force calculation. . . 60
3.50 OpenModelica complete models . . . 61
4.1 Landing gear model. . . 64
4.2 Modify XML file. . . 65
4.3 Java-OpenModelica models interface. . . 66
4.4 Optimization algorithm flow chart. . . 67
4.5 Mass estimation of the BLDC [16]. . . 69
4.6 Pump mass estimation based on the displacement [16]. . . 69
4.7 EMA scaling laws [6]. . . 70
5.1 Actuator anchorage points. . . 74
5.2 Landing gear free body diagram. . . 74
5.3 Ball and roller screws select [17]. . . 75
5.4 Elasto-Gap dynamic. . . 76
5.5 Comparison between linear and non-linear models. . . 78
5.6 Root-locus and poles position of the ˙ϑm/V transfer function. . . 78
5.7 Closed-loop step response . . . 79
5.8 Motor speed of the linear and non-linear models. . . 80
5.9 Location of the closed-loop poles. . . 80
5.10 Closed-loop step response . . . 81
5.11 Aerodynamic force computation. . . 81
5.12 Power flows of the EMA . . . 82
5.13 Power flows of the EHA . . . 83
LIST OF FIGURES LIST OF FIGURES
5.15 Comparison between EMA and EHA electric power consumption. . . 84
5.16 Electric motor torque of both actuators. . . 84
5.17 Force applied by the EMA and the EHA. . . 85
5.18 Applied voltage evolution. . . 85
5.19 Pressure in EHA chamber. . . 86
5.20 Percent error between the reference signal and the electric motor actual speed. . . 86
5.21 Lock force evolution. . . 87
5.22 Current in electric motor phases . . . 88
5.23 Electric power consumption. . . 88
5.24 Electric motor speed . . . 89
5.25 Actuators mass estimation. . . 89
5.26 Initial actuator anchorage points. . . 90
5.27 Optimized actuator anchorage points. . . 91
List of Tables
2.1 Main features of Ball Screw, Recirculating Roller Screw and Planetary
Roller Screw. . . 13
2.2 Landing Gear component drag coefficients [18]. . . 16
3.1 Working principle of the H-bridge switches. . . 33
5.1 Landing gear parameters. . . 74
5.2 Screw parameters [2]. . . 75
5.3 Electric motor parameters. . . 76
5.4 Gearbox parameters. . . 76
5.5 Step response characteristics. . . 79
Nomenclature
Greek Symbols
α Landing gear retraction angle rad
β Effective bulk modulus bar
µ Fluid dynamic viscosity P a · s
ω Electric motor rotational speed rpm
ωn Natural pulsation rad/s
ρ Density kg/m3
τ Pump torque N · m
ϑm Electric motor mechanical angle rad
ζ Damping coefficient −
Roman Symbols
∆P, dp Differential pressure P a
∆S Actuator stroke m
˙m Mass flow rate kg/s
Apiston Cylinder piston area m2
CD Drag coefficient −
CSV Leakage coefficient −
D Aerodynamic drag N
Dpump Pump displacement m3
e, E Electric motor back-electromotive force V
Fact Actuator force module N
g Gravity acceleration m/s2
i Electric motor current A
Kt Electric motor torque constant N · m/A
Kv Electric motor back-electromotive force constant V · s/rad
L Electric motor inductance H
M Landing gear total mass kg
m Mass kg
PAllow Max allowable pressure P a
q Aerodynamic pressure P a
Q, q Flow rate m3/s
R Electric motor resistance Ohm
Re Reynolds number −
Sref Reference surface m2
Te Electric motor torque N · m
TL Load torque N m
V Phase to phase voltage V
vm Actuator average speed m/s
Chapter 1
Introduction
1.1
Motivation
The competition in aviation industry is growing continuously. According to Airbus Global Market Forecast, air traffic is expected to grow at 4.4 % annually until 2037, so that aerospace industry will have to face both economic and environmental issues [19]. For this reason, aircraft manufactures are focusing on new technologies, which can influence overall costs, fuel consumption and reduce emissions. In this context, not only optimizing aerodynamics and engines is important, but also the subsystems design is crucial as it re-duces operation costs by minimizing maintenance and maximizing availability [20, 21]. In order to supply these systems, there are typically three different types of secondary power sources, hydraulic (flight control surface actuation, landing gear extension/retraction and steering, braking, doors), pneumatic (cabin pressurization, air conditioning, de-icing pro-tection) and electric (avionics, cabin, anti-icing, lights, pumps, fans). These three kinds of power sources have their specific features and implementations. In order to achieve the safety requirements, each of them needs a certain level of redundancy. For example, a typical civil transport aircraft normally has up to three on-board centralized hydraulic systems [1]. This strategy results in an increasing waste of power and growing aircraft weight. Moreover, since these power sources are so different from each other, maintenance costs are very high [3]. Furthermore, the engine bleeding affects engine performances [22]. In this context, an attractive solution is represented by the More-Electric Aircraft (MEA) concept, which is based on the progressive elimination of hydraulic, pneumatic and mechanical power sources, replacing them with electric power. This is a key step towards the final goal represented by the All Electric Aircraft (AEA) concept, which shall reduce system weight, complexity, environmental impact, power consumption, logistic and thus acquisition cost [3, 23, 24].
Within the MEA philosophy, a significant effort has been made concerning the ac-tuation technology. Most of the commercial and military aircrafts in service today use centralized hydraulic systems for flight controls, landing gear actuation, brakes, thrust reversals, etc. This system is made of engine driven pumps, which pressurize the fluid used for servo-controlled actuators. It is a time-proven solution that has been matured through decades of aeronautical experience [25]. However, this kind of actuation system is characterized by several drawbacks and limitations [26, 27, 25].
In a centralized hydraulic system, pumps must supply fluid at constant pressure inde-pendently from the actual actuator’s need. Being the pumps driven by the engines, the flow rate is strictly related to engine speed, which depends on the mission phase. This is
1.1 Motivation 1. Introduction
a drawback because the pump must be designed to provide a minimum amount of fluid for the whole speed range. Therefore, for most of the time the pump provides a low flow with respect to its own capacity. This flow is generally not sufficient to dissipate the heat generated by the pump itself. For this reason, it is necessary to ensure a permanent flow inside the pump from the inlet to the outlet with major disadvantages in pump efficiency [1]. One of the crucial drawbacks of having a centralized power generation system, is represented by the hydraulic network needed to provide power to the users, which may be located even 50 meters away from the pumps in large aircraft. This is achieved by fluid-filled pipes, which run throughout the aircraft. It has been estimated that 75 % of hydraulic system mass is due to the distribution network [1].
Moreover, transmitting power by means of mass transfer has several disadvantages. In military aircrafts, synthetic and not environment friendly fluids are used. In com-mercial aircraft, the fluid is less flammable but has a high potential to harm people and environment. Besides, the fluid must operate in specific uncontaminated conditions and temperature ranges, thus, cleaning, purge and contamination controls must be per-formed. Finally, the design of distribution network is subjected to specific constraints such as avoiding critical areas, such as passenger cabin and fire zone [1].
Finally, the metering function, generally performed by a servo valve, causes important power losses. At least 30 % of power is lost at servo valves level [28].
To sum up, the main drawbacks of conventional hydraulic technology are:
– the presence of a heavy distribution network, which increases system complexity and weight;
– low efficiency due to hydraulic losses at pump and metering level;
– the need of an hydraulic fluid, which is harmful to both people and environment; This explains why in recent years, a great effort has been made in the development of Power-by-Wire (PbW) actuators. In contrast to actuators powered by centralized hydraulic system, electric actuators have the advantage of "Power on Demand", consuming power only when needed. Furthermore, as the acronym PbW suggests, they require only electric energy as input, which is provided by cables. This allows removing heavy hydraulic lines, reducing system complexity and maintainability.
Thanks to recent improvement in the electric motor domain and to the increasing maturity of high power electronics, the technology of Electro-Hydrostatic Actuators and Electro-Mechanical Actuators has become sufficiently mature to be introduced in the latest commercial aircraft programs.
Nevertheless, the introduction of new technologies always requires new design paradigms. Indeed, the employment of conventional subsystem solutions, like the centralized hydraulic system for actuation, allows the conceptual design phase of the traditional aircraft not to focus on aircraft subsystems, which are studied mainly in the late preliminary phase. This is made possible by a continuous update of historical databases and the relative absence of interaction between different subsystems. This approach is not valid anymore for the conceptual design of a MEA since historical data is not available and the coupling between subsystems may be important [25]. In addition, the aircraft electrification points out the necessity of implementing smart logics in order to provide the available energy to the different users. The design of these logics requires the power absorption characterization of all aircraft systems [29].
1. Introduction 1.2 Objectives
Due to all these reasons, a great effort is required in order to develop models, which are capable of analyzing aircraft subsystems much earlier than in the traditional design pro-cess. This allows the designer to compare different solutions with only limited knowledge regarding aircraft geometry and other design characteristics [29, 25].
1.2
Objectives
As flight control actuation systems are increasingly becoming electrically powered, there is no more reason to use central hydraulic power to supply landing gear actuation system, since it is the largest short user of hydraulic power on civil transport aircraft [3].
The aim of this work is the development of a model to study landing gear actuation by means of EHA or EMA with particular attention to:
1. Power absorption
2. Electric motor performances
3. Forces transmitted between structural components
4. Losses in the mechanical transmission and hydraulic circuit 5. Influence of different parameters on system performance 6. Transitory dynamic
At the same time, this thesis aims to implement an automatized user-OpenModelica interface using Java language, as shown schematically in Figure 1.1 [30, 31].
Landing gear data Simulation results
Figure 1.1: Schematic representation of the interface implemented in Java Environment. This approach is used to allow the re-use of models and the implementation of op-timization algorithms. By mean of this interface, the user should be able to change the landing gear layout, vary the actuators parameters, adjust control system gains and evaluate the effects without recompiling the model.
Finally, in order to assess the actual usefulness of this study, the models and interface are applied to a practical case study.
1.3
Approach and Methods
The design of landing gear actuation system is particularly complex because it is multi-disciplinary, multi-criteria and a strong coupling between different domains is present [6].
1.4 Outline 1. Introduction
In general, there are two possibilities to tackle this problem [6]. The first one is based on transfer functions or state-space representation of the system. This approach is especially suitable for control design purposes, but simulation results can be approximated, be-cause simplifying assumptions may be necessary to linearize the system. The second way consists of using object-oriented modeling, in which each object (belonging to a specific class) represents a physical component (such as an electrical resistance, a gearbox or a pump) whose behaviour is described by appropriate equations. Each connection between the objects, represents actual physical coupling (for example a power line, a mechanical connection or heat flow). This approach has proven to be very useful for modeling and simulation of complex physical systems [6].
In consequence, starting from the theoretical study of electric actuation technology on one hand, and the analysis of different landing gear retraction mechanism on the other hand, the models of a landing gear, an EHA and an EMA were developed within Modelica Environment. Modelica offers all of the advantages of an object-oriented, multi-domain modeling language [28]:
– It has a multidisciplinary modeling capacity, which means it is able to model com-ponents from different engineering domains.
– It is mainly based on equations instead of assignments. In equations the data flow directions cannot be specified so it can realize the non causal modelling.
– It is object-oriented, which allows the reuse of code.
The development of the user-OpenModelica interface has been realized within Eclipse environment [30].
With regard to the practical case study, once the models of the landing gear and the actuators have been developed, a linearization has been performed and the state space representation has been obtained. At this point, Matlab & Simulink have been used for the actuators control system design and for post processing phase [32].
1.4
Outline
This thesis is divided into the following chapters:
Chapter 2: In this chapter the state of the art and the main research programs on
electric actuation technology are presented. Finally, the basic requirements concerning landing gear retraction mechanism are derived.
Chapter 3: This chapter describes the models developed in OpenModelica of the
landing gear, the EMA, the EHA, the Lock System and the aerodynamic force.
Chapter 4: In this chapter the implementation of the user-models interface is
de-scribed. The classes that allow to quickly modify the models parameters are illustrated. Furthermore, the methods for the actuator mass estimation are presented. Finally, an algorithm for the optimization of the retraction mechanism efficiency is described.
Chapter 5: In this chapter the different models are employed for the study of a
practical case.
Chapter 6: In this chapter conclusions are drawn and avenues for future work are
Chapter 2
Literature Review and Requirements
2.1
State of the Art
12.1.1
Introduction
As stated in the previous chapter, aerospace industry exhibits an inexorable tendency to move towards More Electric solutions. Nevertheless, between a MEA or AEA solution concept and its actual operational employment, several decades will pass. More precisely, the "maturity" of a solution can be identified by the so called Technology Readiness Levels (TRL), which goes from TRL1 (scientific research) to TRL9 (ready to be put into service) [6]. For example, EHAs have been put into service after 30 years from the first concept (1980s) [6]. Therefore, an abrupt switch to Electro-Mechanical Actuator implementation is unlikely to happen. So far, the EHA concept seems to be the best solution for the transit to the All Electric Aircraft, since it combines the advantages of the well-know hydraulic power (damping and low risk of jamming) to the control and wiring advantages of electric system. This process is conveniently represented in Figure 2.1, which visualizes the evolution of actuator technology: firstly, the improvement of actual centralized hydraulic system, secondly the combination of both hydraulic and electrical technologies and finally, the removal of hydraulic technology [1].
This chapter does not focus on the improvements made on the centralized hydraulic system, which ranges from reducing losses at the servo valve level to increasing operating pressure with a consequent increase in power density [1]. Conversely, the most important research programmes concerning EHA and EMA technology are summarized in chrono-logical order, with particular attention on landing gear actuation. Then, the state of the art of the actuators main components is presented.
2.1.2
Research Programmes on EHA
EHA with variable displacement pump (EHA-VD) were the first to be developed, when motor control electronic was not mature enough to perform current switching. In this case, the electric motor (generally Alternating Current induction motor) is directly connected to the AC supply [21]. A movable mechanical component as a swash plate is employed to perform variation in pump displacement. This type of actuator has been put into service
1Most information on the electric motors and EMA mechanical transmission state of the art is gathered from notes of Impianti Aeronautici II lessons, by Prof. Ing. Roberto Galatolo full professor at University of Pisa.
2.1 State of the Art 2. Literature Review and Requirements
Figure 3.2. Evolution of actuation (top to bottom, from all hydraulic to all electric)
Figure 2.1: Actuator evolution from centralized hydraulic circuit to Elecetro-Mechanical actuator [1].
in the early 50s on the bomber Avro Vulcan, in which 10 Powered Flying Control Units were employed for flight controls [1].
The IAPTM (Integrated Actuator Package) is an implementation of the EHA-VD
con-cept. They were developed by Lucas in the 80s in order to improve survivability of military aircraft, in which the hydraulic network increases vulnerability to enemy fire [1]. The operation diagram is presented in Figure 2.2, in which the main characteristics of this actuator are highlighted: 1) overload protection 2) cavitation protection 3) declutching 4) second generator to supply the control unit 5) displacement control unit.
The studies on this actuator identify the following issues: low efficiency (mainly due to the asynchronous motor), difficulty to predict heat generation and need of proper fluid pressurization [1]. EHA-VD technology is not used at landing gear level, as flow controllability does not make much sense for this purpose [3].
The second concept of EHA is the Electro-Hydrostatic actuator with Fixed Displace-ment Pump (EHA-FD), in which a speed controlled electric motor drives a fixed displace-ment pump that feeds a conventional hydraulic cylinder. Change in direction is allowed by a bidirectional motor. This solution has proven to be simpler (in terms of maintenance and manufacturing) than the EHA-VD [21].
In the USA, EHA-FD have been developed first for military application in order to increase survivability. In the 90s, a tandem aileron actuated by a EHA-FD has been put
2. Literature Review and Requirements 2.1 State of the Art
Figure 5.3. Power architecture of an IAP actuator
Figure 2.2: IAP working scheme [1].
into service on the HTTB C-130 2 [1]. Electrically Powered Actuation Device (EPAD),
Electric Actuation and Control System and Fly-by-Light Advanced Systems Hardware are some of the most important research programs that have been started since then [1]. In particular EPAD, a partnership between the US Air Force (USAF), Navy, and NASA, has led to flight-tested EHA-FP for aileron control on the F-18 [33]. Several ground tests pointed out the feasibility of EHA-FD implementation on military aircrafts. In October 2000, F-16 was the first aircraft that used exclusively PbW actuators.
Likewise, starting from the late 80s, the EHA-FD technology was investigated in Eu-rope, but not for military applications. The objective was to eliminate several hydraulic networks of the centralized hydraulic system on commercial aircrafts. The studies have fo-cused on the aileron actuation, since it is permanently loaded. Different research programs were carried out. Among them, All Electric Flight Control Actuation (ELAC) program goal was to test the EHA-FD, EHA-VD and EMA for comparison purposes. Within the Commandes de Vol Futures – Future Flight Controls program by Lucas Air Equipment, an Electro-Backup-Hydrostatic Actuator (EBHA), in which the servo hydraulic actuator and EHA-FD concepts are combined, was flight-tested on the Airbus A320. In 1997, Liebherr flight tested an EHA-FD on the Airbus A321 within the Electrically Powered Integrated Control Actuator program. A more recent program, the Power Optimized Air-craft (POA)(2001-2007), aimed to the standardization of the electro-hydrostatic modules.
EHA-FD in Landing Gear Actuation
In general, actuators on landing gear perform several functions: – Doors actuation
– Up and Down Lock actuation – Extension/Retraction
– Steering and Braking
For most of these function, a large amount of power is required in a very short time and the centralized hydraulic system has demonstrated to be a well-proven solution. However, the need for hydraulic supply from the centralized system leads to a complex and heavy
2.1 State of the Art 2. Literature Review and Requirements
distribution network, especially for the nose landing gear, which is placed even 40 meters from the hydraulic pumps [1].
So far, PbW actuation for landing gear has not been investigated as deeply as for flight controls. In 1992, an EHA concept for landing gear extraction/retraction was ground tested in Japan, although this demonstrator did not fulfill reliability and mass constraints. Within the POA program, a fully nose landing gear actuated by EHA-FD was tested. The objective was to use sequential power to fulfill the Landing Gear Extension-Retraction System (LGERS) function. In 2006, the More Electrical LANding gear sYstem program investigated the possibility of steering by means of EHA (see Figure 2.3).
156 Aerospace Actuators 2
Figure 5.7. LGERS system with EHA-FD within the POA program [GRE 04]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip Launched in 2006, the MELANY (More Electrical LANding gear sYstem) program started by exploring the landing gear steering by EHA. Figure 5.8 presents the electro-hydrostatic module and its integration on a nose landing gear of the Airbus A320.
Figure 5.8. Landing gear steering by EHA as part of the MELANY project, extracted from [LIE 12]
Figure 2.3: Landing Gear Steering Actuator [1].
The objective of the Technology for Hydraulic-Electric Retraction Mechanism Actua-tion Equipment (THERMAE II) program was to demonstrate that a LGERS applicaActua-tion for the main landing gear of the A320 had reached TRL5. The architecture was made of three actuators for doors opening, extension/retraction operation and down-locking (see Figure 2.4). Two motor-pump unit provided hydraulic power generation, while a single control unit performed motor control, monitoring and valve opening control. In addition to current and motor speed loop, the system was provided with a pressure loop to detect the piston approaching the end stop. Hence, no position sensors are required [1].
158 Aerospace Actuators 2
pressure loop to transient changes of mode. Although mass and reliability balances may be negative at the level of the equipment compared to a conventional solution, the developed solution was expected to generate overall potential advantages for the aircraft.
158 Aerospace Actuators 2
pressure loop to transient changes of mode. Although mass and reliability balances may be negative at the level of the equipment compared to a conventional solution, the developed solution was expected to generate overall potential advantages for the aircraft.
Figure 5.9. LGERS system with EHA-FD within THERMAE II program. Upper image: power architecture [ELL 16], lower image: overall view
Figure 2.4: EHA architecture and overall view of the EHA-FD of THERMAE II program [1].
2. Literature Review and Requirements 2.1 State of the Art
2.1.3
Research Programmes on EMA
The concept of EMA is based on the removal of hydraulic fluid as a vector for power transmission. In general, direct drive by an electric motor is not possible since the low torque/high speed of the motor output is incompatible with the high force and low speed needed. Consequently, a gearbox is generally connected to the electric motor. If a linear movement is desired, a screw can perform the requested reduction without gearboxes (Direct Drive EMA). However, a gearbox is often used between the motor and the nut screw to accomplish specific requirements (Geared EMA) [1].
The absence of hydraulic fluid leads to several issues, which need to be tackled [1]: (a) Necessity to damp loads
(b) Snubbing function and declutching are no longer possible (c) Hydraulic fluid was used for heat transferring functions (d) Increasing in reflected inertia
(e) Gear backlash can lead to oscillations and control system instability (f) Service life depends on wear of mobile parts such as ball or roller screws
(g) Jamming, which is critical for flight control and landing gear actuation. May be tolerated for other application as the trim of the horizontal stabilizer
Numerous research programmes on the development and implementation of EMA have been launched for space applications, however these will not be covered in this thesis.
Regarding aircraft primary flight controls, Advanced Electro Mechanical System was one of the first programs, which led to the flight testing of EMA. The actuator was implemented on the Lockheed C-141 for aileron control. Program results highlighted the importance of mechanical damping [1].
In the 90s, the American program Electrically Powered Actuation Device compared three actuation technologies on the F-18 aileron control [33]. The first actuator was a Smart Actuator, in which the failure detection logic was implemented locally and not centrally (by the Flight Control Computer) as it is usually done. The second actuator was an EHA and the third an EMA. As for the latter, the tests conducted within the program aimed to compare the performance of this actuator with that of the standard hydraulic actuator under actual load and environmental conditions. The main features of the EMA are briefly summarized in the following. A Power Conversion Unit was used to transform the 115 V AC voltage into 135 V Direct current (DC). The aircraft has also been equipped with all the necessary devices to transmit data to the ground in real time. The EMA consisted of two 3-phases Brushless motors connected to a single ball screw in velocity-summing configuration. The stroke was 4.125 in, the maximum load was 13.200 lb and the weight was about 26 lb. An anti-rotation device was integrated. The control unit provided closed-loop position control with three nested loops on the current, speed and position. Besides, the Power Control and Monitor Electronics (PCME) performed monitoring operations. Phase switching was performed by MOS-controlled thyristors. Pulse Width Modulation (PWM) was used to modulate power. The actuator was also equipped with an Interface Box. This box was used to interface the PCME with the FCC. In particular, the box collected data from the PCME and sent it to the instrumentation
2.1 State of the Art 2. Literature Review and Requirements
system through the MIL-STD- 1553B data bus. The test lasted 25 flight hours, during which Mach 1.54 and 6-g of vertical acceleration were reached. The performance in terms of response to commands of the EMA-driven actuator was compared with that of the standard actuator mounted on the aileron of the opposite wing. The results showed similar performance for the two actuators, apart for a better dynamic response of the electric actuator. The most significant problem highlighted by the program was thermal performance. It was expected that the most challenging thermal conditions would be encountered during severe manoeuvres imposed on the aircraft. Though, the critical condition was to keep a load in place while using the ailerons as flaps. Other lessons-learned were the importance of ensuring screw rotation by means of no-rotating devices and the importance of mechanical end-stops [33].
Meanwhile in Europe, the same problems were concluded during the ELAC program. The More Open Electrical Technologies project, related to the aileron actuation of the Air-bus A320, focused on the calculation of actuator service life in order to develop monitoring algorithms. The COmmande de Vol Avec Distribution de l’Intelligence et Intégration du Système: Flight Control with Distributed Intelligence and Systems Integration program aimed to prove that EMA technology had reached TRL6. Between 2011 and 2016, the "Actuation 2015" project focused on standardization of EMA technology.
Nevertheless, EMAs are still not considered matured enough to be used for primary flight controls. This does not apply for secondary flight controls, where jamming is con-sidered an "acceptable" failure mode. This explains why EMA technology has been put into service for Airbus A380 slats and the trim of horizontal stabilizer on Airbus A350. Within the Advanced Flight Control program, EMAs where tested for spoiler actuation [1].
EMA in Landing Gear Actuation
Focusing on the landing gear, EMAs perform different functions.
In the braking system, the EMA’s task is to produce the axial force on the back of the braking disks. They are successfully employed on the Boeing 747 electrical braking system [1]. Regarding the extension/retraction system, the mechanical jamming rep-resents the main difficulty of employing EMA for this function, since it reprep-resents an unacceptable failure mode. As a matter of fact, in case of total loss of supply power, an emergency extension of the landing gear has to be performed [5]. In a conventional hydraulic cylinder, this is not hard to achieve, since damping can be performed hydrauli-cally by placing a throttle valve between the cylinder chambers. This is no longer possible if EMAs are employed [34]. This is why the most important related research programs have focused on this problem. This is the case of the ELGEAR, Concept Innovant de Système d’Actionnement de Commandes de vol secondaires et de Servitude: Innovative actuation system for secondary flight controls and utilities and AiRcrafts Main LandIng Gear acTuation projects. In the last one, an internal rod was placed inside the screw and connected to the load. In case of jamming, this internal rod is separated from the screw (see Figure 2.5)[1].
Also, in steering it is necessary to assure self-alignment in case of actuation failure. Besides, damping must be guaranteed in order to assure for shimmy. Despite of the extension/retraction function, steering actuators must perform a precise position control. Investigations have been made within the Distributed and Redundant Electro mechanical nose-gear Steering System [1], which objective was the development of a PbW steering system with high reliability for the Airbus A320 [1].
2. Literature Review and Requirements 2.1 State of the Art
Electro-mechanical Actuators 189
Figure 6.7. Examples of EMA extension/retraction demonstrators
In 2005, the DRESS (Distributed and Redundant Electro mechanical nose gear Steering System) project aimed to develop a PbW steering system with high reliability ( λ < 10−9/FH), which authorized zero visibility landing
(category IIIC) for an A320 aircraft [IOR 10]. Among the evaluated power architectures, the developed and ground tested solution implemented two electromechanical paths (remote motor control electronics, PMSM
Figure 2.5: EMA demonstrator [1].
At this point, one might question what is the advantage of switching to EMA since this technology still has several unresolved issues. Employing EMAs is not only pursued because it is a crucial step towards the AEA concept but also because the absence of hydraulic fluid allows to reduce complexity and thus maintenance cost. Furthermore, stiffness is increased because of the lack of fluid in the load path [23, 33].
2.1.4
Electric Motors
In general, electric motors may be divided into three categories: Induction Motors (the so called Asynchronous Motor), Reluctance Motors and Synchronous Motors. On the one hand, Induction Motors are inexpensive and they do not require excessive maintenance. On the other hand, low efficiency is one of the main drawbacks. For this reason, they are not used for aerospace applications, where high toque and high dynamic performances are required [13].
Thanks to recent research achievements, the Permanent Magnet Synchronous Motor (PMSM) has become a reference technology for aircraft actuation. PMSMs are charac-terized by high efficiency, fast dynamic response, low torque ripple and improved thermal behaviour, which results in major torque capability. The principal drawback is the need of an expensive sensors technology in order to track shaft position and power electronics to perform phase isolation [13]. The current-torque relationship is quite linear, which is ben-eficial to control design. These motors are often called Brushless Motor (BLM) because of the similarities with DC motors. In fact, both motor working principles are based on phase commutation. However, in the DC motor commutation is performed mechanically by brushes, while in the PMSM commutation is conducted electrically (that is why they are called "brushless"). Synchronous refers to the angular orientation between the static magnetic field axis and the rotatory magnetic field axis, which is kept constant during motion. More in general, BLMs may be classified into two categories based on voltage modulation and control methods:
– Permanent Magnet Synchronous Motor (PMSM) characterized by a sinusoidal-wave back-electromotive force (emf)
– Brushless Direct-Current (BLDC) characterized by a trapezoidal-wave back-emf The first one, is characterized by low torque ripple, high efficiency and low noise but rotor position must be known continuously. In contrast, trapezoidal modulation allows simpler switching logic and this increases reliability but torque ripple is significant. Both technologies are characterized by [13]:
– high power density (2.5 kW/dm3)
2.1 State of the Art 2. Literature Review and Requirements
– high efficiency (0.9) – fast dynamic response
Switched Reluctance Motor (SRM) is cheap, compact, and does not need a complex power electronic, which means lower overall cost. The main disadvantages are a quadratic current-torque relationship and high torque ripple. The principal characteristic of this motor type is the absence of magnets on the rotor. With regards to the stator, it is similar to the PMSMs. Phase current generates a magnetic flux and a magnet torque is generated in order to reduce stator-rotor reluctance.
2.1.5
EMA Mechanical Transmission
EMA employment opens to the possibility to have a torque/rotational output instead of a linear one, typical of a hydraulic cylinder. Nevertheless, this implies to reconsider airframe design and question a solid know-how. Due to this reason most of the research programs have focused on linear EMAs. In EMA mechanical transmission an intermediary reducer may or may not be required. The addition and choice of the reduction ratio depend on the following considerations [1]:
– Motor mass: it depends on the torque that must be produced so it may be interesting to use a reducer but its additional mass must be taken into account. Moreover, using a satellite roller screw allows to increase reduction ratio so that the intermediate reducer can be removed. However, a satellite roller has a lower efficiency than a ball screw so that a greater torque could be required.
– Reflected inertia: the equivalent load mass is a squared function of the reducer ratio. The first consequence, is the increasing kinetic energy that has to be absorbed by the mechanical end stops:
Meq = K2J (2.1)
K= 1
2MeqVapp2 (2.2)
Where Meqis the load equivalent mass, K is the total reduction ratio, J is the motor
inertia K is the load kinetic energy and Vapp is the load approaching speed.
An other consequence of increasing reflected inertia is the reduction of natural sys-tem frequencies.
– Backlash: the employment of a reducer introduces inevitably a functional backlash. This has important consequences on control system stability. It is possible to remove backlash by introducing preload but this results in service life reduction since the two faces or threads of the reducer are contemporaneously in contact.
Regarding the gearbox technology, harmonic gear, cycloidal, or planetary gear reducers are all valid options because of their compact structures and high efficiency. In particular, harmonic and cycloidal reducers allow to reach high reduction ratios, which results in significant motor mass reduction [6].
With regard to screw technology, the most commonly employed solutions are ball screw and roller screws. More precisely, research programs started to investigate roller screw solutions in order to overcome the ball screw’s jamming problem [13]. Roller screw
2. Literature Review and Requirements 2.1 State of the Art
can reach higher rotational speeds thanks to balls elimination and stand higher loads as a consequence of a greater pressure transmission area, as shown in Figure 2.6:
Roller screws
conceptual design
advantage over ball
screws
With planetary roller screws, the application load is transmitted from the nut to the shaft through the barrel–shaped surfaces of the rollers. The number of contacts and the total surface area of the contacts between the shaft, the rollers and the nut are substan-tially increased compared to the ball screw design, resulting in larger dynamic and static load carrying capacities († fig. 1).
The absence of recirculating element embodies the fundamental conceptual ad-vantage of the planetary roller screws. This feature eliminates the main failure mode of ball screws, that is linked to the recirculation of the balls. Indeed, recirculating heavily loaded balls induces alternate stresses on the balls and shock loads arising from the change of trajectory.
In addition, satellite rollers never come into contact with each other. This is a signifi-cant advantage of this product over most ball screw designs. Balls come into contact with each other in most ball screw designs, generating friction and adding a potential failure mode to the ball screw concept.
With recirculating roller screws, the ap-plication load transfers from the shaft to the nut through a set of grooved rollers. This design permits very small leads while offer-ing high load carryoffer-ing capacity and axial stiffness. This mechanical advantage mini-mizes input torque and increases application resolution and performance. They can sim-plify a complete transmission and improve its rigidity. They are often used in applica-tions involving advancing technology where reliable optimum performance is essential.
Roller screw concept Advantages over a ball screw User benefits Large number of contact
points Satellite rollers
Planetary roller screws with small lead (down to 2,00 mm)
Evenly spaced planetary or recirculating rollers Recirculating roller screws with small lead down to 1 mm
High load carrying capacity and up to 10 times longer service life
Roller screw rotational speeds up to 50% higher than a ball screw with similar carrying capacity
Roller screw acceleration up to 3 times higher
Absence of recirculation eliminating a significant failure mode
High load carrying capacity compared to ball screws with small lead which are designed with small diameter balls which have low load carring capacity Good operation in applications with changes of direction, stable friction torque
High load carrying capacity, high axial stiffness that cannot be obtained with a ball screw of similar lead and diameter Very small imput torque
Lower total cost of ownership (TCO) Increased speed of operations Higher productivity Higher degree of reliability High load carrying capacity combined with positioning accuracy and reduced torque requirements Low noise
High degree of reliability High resolution, high stiffness, long service life, robustness
Fig. 1 Comparison of ball screws and roller screws contact area
18 Figure 2.6: Comparison between roller and ball screw contact area [2].
For all these reasons a roller screw will most likely be used for landing gear actuation. Roller screws may be further divided into planetary roller screw and recirculating roller screw. In the latter one, the circumferential grooves of the rollers reproduce the ball-type contact whereas the planetary roller screw emulates the kinematic of a planetary gearbox. As the mechanical contact is very different in these three typologies of screw technology, each of them has different features, as pointed out in Table 2.1 [13].
Table 2.1: Main features of Ball Screw, Recirculating Roller Screw and Planetary Roller Screw.
Ball Screw Recirculating Roller
Screw Planetary Roller Screw
Low friction High friction High friction
High pitch Low pitch Low pitch
Low load capacity High load capacity High load capacity
Lighter than roller screw No need of re-circulation
Jamming problem due to ball re-circulation
Concerning the motor-nut screw layout, the most common solutions are presented in Figure 2.7.
2.1 State of the Art 2. Literature Review and Requirements
Electro-mechanical Actuators 201
Figure 6.14. Motor/nut–screw integration in an EMA. For a color
version of this figure, see www.iste.co.uk/mare/aerospace2.zip
Figure 6.15. Nut–screw layout in an EMA, according to [KAR 06]. For
a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip
Figure 2.7: Different Motor-Screw layouts [1].
The third solution has few applications in aeronautics, since it requests conical gears, which have more constrains than spur gears [1]. The last two solutions allow overall length reduction and are therefore more interesting. However, they imply the use of an intermediary gear.
In a nut-screw system the nut can be rotating and the screw translating or vice versa. Each configuration has specific constrains on the motor-gear layout. The rotating part must be ensured against translation, while actions must be taken to prevent the translating part from rotating . The axial reaction force is the equivalent of the pressure in a cylinder chamber. What is new, is the pick-up torque, which must be balanced by a torque link outside or inside the body.
Regarding declutching or locking functions, these are performed by brakes and elec-tromagnetic clutches. The objective is to produce a fail safe redundancy architecture [1].
Moreover, damping at end stop must be considered. It is necessary to limit kinetic energy as the actuator reaches the end stop. In landing gear extension/retraction, this function is often implemented on a control level, which means the controller forces decel-eration when the nut (or the screw, depending on which part is translating) approaches the end stop. However, this requires sensors able to detect approaching. Other solutions are a torque limiter or elastic elements [1].
Finally, overload protection has to be implemented: this function may be achieved by introducing current saturation.
With regard to control implementation, for an end-stop to end-stop function, a typical control sequence is (extension):
(a) Unlock
(b) Acceleration phase to motor nominal speed followed by a constant speed phase (c) Deceleration phase triggered by a sensor position, which informs of approaching to
the end-stop
(d) Approach at low speed. Current (that means torque) may increase in order to engage the lock
2. Literature Review and Requirements 2.2 Design Requirements
2.1.6
EHA Hydraulic Circuit
The main element of the EHA in addition to the electric motor is the pump. EHA pumps operate in more severe conditions than centralized hydraulic pumps. The last ones, operate in a single power quadrant, in a relatively narrow range of speeds and at quasi-constant differential pressure. For this reason, they can be designed and optimized to work at a specific design point. On the contrary, EHA pumps operate in four power quadrants. Speed can reach 16000 rpm in both directions and differential pressure may exceed 350 bar [1].
For this reason, EHA pump design is a hard task, since different operating conditions must be taken into account. Besides, mechanical and volumetric efficiencies have a direct impact on the torque that has to be produced by the motor. The challenge is not so much to minimize the mass for an operating point as to minimize losses for mean operating conditions [1].
Concerning pump technology, so far axial piston pumps represent the best solution. They can operate at very high pressures and in a wide range of speeds. However, service life of these pumps is still not long enough to allow EHAs to be employed as front-line actuators for primary flight controls [1].
With regard to the other elements of EHA hydraulic system, it must be borne in mind that this is a closed circuit. Hence, any gas presence greatly affects the actuator stiffness and thus, control stability. For this reason, fluid filling procedures must be carried out very carefully. Further, a circuit pressurization is very important to contrast cavitation. A possible solution is increasing accumulator charging pressure. However, this puts seals under permanent stress. The accumulator is also in charge of leakage compensation. Therefore, this element should be sized in order to compensate the external leakages that occur during service life. Nowadays, EHAs are connected to the central hydraulic network so that accumulators are sized only to compensate geometrical effects (differences in piston areas) and thermal expansion [1]. Finally, several valves are available to allow the circuit to work properly, such as bypass valves and check-valves.
2.2
Design Requirements
2.2.1
Introduction
Landing gear actuation system design must meet several requirements. High reliability, minimum weight, good maintainability and low cost are generally demanded [3]. Being the safety requirements the most important ones, a certain level of redundancy is needed. In more electric power actuation, redundancy can be obtained on electrical power source level and on actuator level. The first one should not be a problem in a MEA. Regarding actuator level redundancy, actuators are typically non-redundant because of space constraints. This means one landing gear leg is moved by one single actuator [3]. Redundancy applies to actuator power conversion devices such as hydraulic pumps and electric motors (example of these are the motors in torque or velocity summing configuration). The redundancy level study is a complex aircraft design activity and it is not covered in this thesis.
2.2.2
Load Requirements
Focusing on extension/retraction requirements, the actuation system should fulfill the following functions [34]:
2.2 Design Requirements 2. Literature Review and Requirements
– Extension/retraction in normal mode – Free-fall extension in emergency mode – Uplock
– Downlock
During extraction/retraction phase, several loads act on the landing gear. Within the thesis, the following static and dynamic loads were taken into account:
– Weight
– g-Loads due to aircraft manoeuvres – Aerodynamic
– Friction
Among them, loads due to gravity and aerodynamic drag are the most significant during landing gear actuation. Regarding dynamic loads, usually the assumption is made that the aircraft performs a 2-g turn [18]. The best way to evaluate aerodynamic loads is through test data for similar gear arrangements. If this data is not available, for initial analysis landing gear drag can be estimated by summing the contribution of its main components, using data of Table 2.2, in which D indicates the resistance force, q the dynamic pressure and Sref the frontal area.
Table 2.2: Landing Gear component drag coefficients [18].
D/q Sref
Regular wheel and tire 0.25
Second wheel and tire in tandem 0.15
Streamlined wheel and tire 0.18
Wheel and tire with fairing 0.13
Streamline strut 0.05
Round strut or wire 0.30
Flat spring gear leg 1.40
Fork, bogey, irregular fitting 1-1.4
Multiplying these coefficients by the dynamic pressure and frontal area leads to re-sistance force of each component. In first analysis, it is possible to take into account the resistance due to the mutual interference by multiplying the result by 1.2 [18]. With regard to the airspeed, the structure shall be designed to withstand aerodynamic loads up to 1.6VS1, in which VS1 is the stalling or minimum speed, in knots, at which the aircraft
is controllable with engines idling [5]. In addition, the loads produced by a crosswind must be taken into account. The speed of this wind is equal to the highest value be-tween 20knots and 0.2VSRO, in which VSRO is the stall speed in take-off configuration
2. Literature Review and Requirements 2.2 Design Requirements
2.2.3
Time Actuation Requirements
Retraction time is generally more significant than extension time, for two reasons [3]: (a) During climb, the overall drag of the aircraft must be minimized and the contribution
of the landing gear is extremely high.
(b) During retraction, the actuation system must produce higher power than during extraction phase.
According to the "European Certification Specifications for Large Aeroplanes" (CS-25) of theEuropean Union Aviation Safety Agency (EASA), actuation time limits are not specified [34].
It is possible to determine the retraction time by considering data available for similar aircraft. The following data was obtained from actual flight tests carried out on different Boeing aircraft. The landing gear was retracted 3 s after the lift-off manoeuvre. In addition, in some cases one pump was switched off to obtain more cautious results. As pointed out in Figure 2.8, the results show an almost linear trend of the retraction time with the Maximum Take-Off Weight (MTOW) [3].
Chapter 3
Landing Gear Actuation System
3.1
Landing Gear Modelling
3.1.1
Introduction
The purpose of this section is to make the reader familiar with landing gear terminology. What is a landing gear? Conway defines it as “the essential intermediary between the aeroplane and the catastrophe” [5]. It represents one of the most critical subsystem of the aircraft and its configuration goes hand in hand with the aircraft conceptual design because of its impact on the aircraft structure itself [35].
By shifting to a more electric actuation, the nose gear and the main gear actuation systems will become independent so that the two systems can be studied separately [3]. This thesis focuses on the main landing gear actuation system.
The main gear supports the weight of the aircraft during ground operation (taxi, take-off roll, landing roll and steering), absorbs and dissipates landing kinetic energy reducing the loads transmitted to the airframe, contributes to breaking capability and protect the ground surface [5].
In order to absolve all of these functions, it includes several structural components and systems depending on the landing gear type (refer to Figure 3.1 and Figure 3.2).
The main component is represented by the shock strut, which objective is transforming the energy of landing impact into heat. There are many types of shock absorbers. Among them, the pneumatic/hydraulic shock strut is one of the most employed on large aircraft. It consists of two telescopic cylinders, in which the upper one is fixed and bound to the aircraft while the lower one is free to slide inside the upper cylinder. Two chambers are thus formed. The lower one is filled with liquid while the upper one is filled with compressed air or nitrogen. When the cylinder undergoes compression, an orifice between the two cylinders provides for the passage of fluid between the chambers. In general, this flow is not constant but is controlled by a metering pin. The hydraulic flow forced to pass trough the orifice causes heat generation [4].
Most struts terminate with an axle at the bottom of the lower cylinder to allow wheel installation (see Figure 3.2a).
As shown in Figure 3.2b, struts are also equipped with torque links in order to keep the piston and wheels aligned. One end of the link is fixed to the upper cylinder while the other end is connected to the lower cylinder to prevent it from rotation. It also allows to hold the piston in the upper cylinder when the strut is extended, for example after take off [4].
