Alternative Flight Control in Civil Aviation

Home , Flight control surfaces, Thrust reversal

D. Geleyns Technische Universiteit Delft

ALTERNATIVE FLIGHT CONTROLIN CIVIL AVIATION

D. Geleyns

in partial fulﬁllment of the requirements for the degree of

Master of Science in Aerospace Engineering

at the Delft University of Technology, June 15, 2016

Supervisor: Dr. ir. R. Vos Thesis committee: Prof. dr. ir. L. L. M. Veldhuis, TU Delft Dr. ir. C. C. de Visser, TU Delft

An electronic version of this thesis is available at http://repository.tudelft.nl/. Thesis number: 080#16#MT#FPP Equivalent word count: 22639

ABSTRACT

This report contains a study about the possible beneﬁts of changing the way large airliners are controlled in ﬂight. The goal is to save fuel by decreasing the structural weight of the aircraft. Over the last decades the conventional primary control surfaces in civil aviation have not deviated from elevators, ailerons/spoilers and rudder. Eliminating one of these by compensating with others could hold revolutionary advantages.

The research started out by finding alternative ways to control the aircraft around its different axes. This performance analysis was done with a Boeing 747-200 model in the RECOVER Benchmark tool with slight modifications. It was found that yaw control offered no decent alternatives besides the rudder. Pitch control can surprisingly be done with outboard ailerons, although it barely meets the requirements. This finding is attributed to the large sweep angle of the Boeing 747. Roll control offered the most alternatives due to many redundant control surfaces on the wings. Further analysis of unconventional roll control was split into four cases. The first case, flying without any ailerons, was determined to be unable to meet the legal requirements. However, the second case in which two spoiler surfaces were added to the existing six, was found to be satisfactory in all phases of flight. The efficiency of controlling the aircraft with spoilers instead of ailerons was addressed and a considerable difference was found but in the overall flight, this is not an issue. The third case was more moderate as it only removes the inboard ailerons. This case had no issues to meet requirements. The final case is entirely different: flying without spoilers. All roll requirements were met, but only after the outboard aileron was enlarged to within structural bounds. The larger issue with this case was the lack of airbraking features. The second phase of the project focused on evaluating the snowball effect in aircraft design caused by the small changes in the roll control mechanisms. First the weight of the removed components was estimated through a hydraulics model. This showed results ranging from 7% to 20% in subsystem weight decrease, depending on the case. The next step was to run these updated weights through a preliminary design phase. To facilitate the process the Aircraft Design Initiator was used. Two aircraft were reproduced and modified to assess the impact on different scales as well: the Boe- ing 747-200 and the Boeing 767-200, which is approximately half the size. It was found that the fuel percentage that could be saved for each case are 0.50%, 0.36%, 0.18% and 0.45% in respective order for the Boeing 747. Interestingly, these percentages are nearly identical to those of the Boeing 767. The impact on the operating empty weights does differ slightly, where the Boeing 747 loses more weight than the 767, percentage-wise.

Being that the fuel saving on a standard harmonic mission proﬁle is less than 0.50% for any of the cases of alternative roll control, the results are not revolutionary as they are now. The proﬁtability of switching to these new control methods is debatable.

iii

PREFACE

This thesis is the final part of my Master program in Flight performance & Propulsion at the Faculty of Aerospace Engineering of Delft University of Technology. It marks the end of my life as an engineering student and opens up a new chapter. I am proud and satisfied with what I have achieved over the past years in Delft and especially with this final milestone. I would like to thank my supervisors, Dr. ir. Roelof Vos and Dr. ir. Coen De Visser, for their guidance, feedback and advice during this research project. I also thank Prof. dr. ir. Leo Veldhuis for presiding my graduation committee. Lastly, I would like to express my gratitude towards my parents Luc and Hilde, my partner Ine and my housemate Toon. They were on the front row during stressful outbursts and they were listeners and contributors when I was brainstorming about the project.

D. Geleyns Hasselt, Belgium June 2016

CONTENTS

List of Figures ix

List of Tables xi

1 Introduction 1 1.1 Context...... 1 1.2 Research question and thesis goals...... 1 1.3 Boeing 747 Flight Control System...... 2 1.4 Report structure...... 3

2 Unconventional Flight Control5 2.1 Unconventional control possibilities...... 5 2.1.1 Roll control...... 5 2.1.2 Pitch control...... 6 2.1.3 Yaw control...... 6 2.2 Analysis of new control opportunities...... 7 2.2.1 RECOVER Benchmark...... 7 2.2.2 Pitch control...... 9 2.2.3 Roll control...... 12

3 Flight performance impact of fewer control surfaces 17 3.1 Case 1: Fewer ailerons...... 17 3.1.1 Regulations...... 17 3.1.2 Landing...... 18 3.1.3 Cruise...... 19 3.1.4 Modiﬁcations...... 20 3.1.5 Additional remarks...... 21 3.2 Case 2: No spoilers...... 22 3.2.1 Regulations...... 22 3.2.2 Landing...... 23 3.2.3 Cruise...... 24 3.2.4 Modiﬁcations...... 25 3.2.5 Additional remarks...... 28

4 Weight Estimation 31 4.1 Hydraulic system model...... 31 4.1.1 Veriﬁcation...... 34 4.2 Aircraft Design Initiator...... 35 4.3 Results and discussion...... 36 4.3.1 Veriﬁcation...... 40

5 Other aircraft 41 5.1 Airbus versus Boeing...... 41 5.2 Boeing 767-200...... 42 5.2.1 Results and discussion...... 43

vii 6 Conclusion 47

7 Recommendations 49

Appendices 51

A hydraulics.m 53

Bibliography 55

viii LISTOF FIGURES

1.1 Control surface layout on a Boeing 747. [1]...... 2

2.1 General Simulink structure of the RECOVER Benchmark...... 8 2.2 Typical command window prompt of RECOVER Benchmark...... 9 2.3 Modifications in RECOVER Benchmark to control the elevator surfaces individually..9 2.4 Altitude and Alpha plots with outboard aileron input during final approach: full upward deflection...... 11 2.5 Altitude and Alpha plots with normal pilot input during final approach: -6 degrees on control column (50%)...... 11 2.6 Altitude and Alpha plots without any control input during final approach...... 12 2.7 Result of a max input roll maneuver following EASA standards...... 14 2.8 Roll response comparison between scenarios with different surfaces at maximum deflection...... 15

3.1 Maximum achievable roll maneuver comparison in landing setup for the aileron case. 19 3.2 Maximum achievable roll maneuver comparison in cruise setup for the aileron case.. 20 3.3 Wing layout of the Boeing 747-200...... 21 3.4 Effectiveness of adding spoilers towards the wingtip...... 21 3.5 Efficiency comparison between aileron roll and spoiler roll maneuvers...... 22 3.6 Maximum achievable roll maneuver comparison in regulation setup for the spoilers case 23 3.7 Maximum achievable roll maneuver comparison in landing setup for the spoilers case. 24 3.8 Maximum achievable roll maneuver comparison in cruise setup for the spoilers case.. 24 3.9 Maximum achievable roll maneuver comparison in regulation setup with addition of rudder deflection...... 26 3.10 Sideslip response from the aircraft’s rudder deflection...... 26 3.11 Expansion of the outboard aileron on a Boeing 747-200...... 27 3.12 Cascade thrust reverser on a Boeing 747-400...... 29

4.1 Flight control hydraulic system on a Boeing 747-100. [1]...... 33 4.2 Flow chart of the Aircraft Design Initiator [2]...... 35 4.3 Maximum take-off mass vs. hydraulics mass for a Boeing 747-200 on a harmonic mission profile...... 37 4.4 Operating empty mass vs. hydraulics mass for a Boeing 747-200 on a harmonic mission profile...... 37 4.5 Fuel mass vs. hydraulics mass for a Boeing 747-200 on a harmonic mission profile... 38

5.1 Maximum take-off mass vs. hydraulics mass for a Boeing 767-200 on a harmonic mission profile...... 43 5.2 Operating empty mass vs. hydraulics mass for a Boeing 767-200 on a harmonic mission profile...... 44 5.3 Fuel mass vs. hydraulics mass for a Boeing 767-200 on a harmonic mission profile... 44

LISTOF TABLES

1.1 Failure probability of individual control systems for a Boeing 747. [1]...... 3

2.1 Unconventional control possibilities around aircraft axes...... 5 2.2 Wing quarter-chord sweep angles for common jet airliners...... 12 2.3 Roll rate comparison between scenarios with different surfaces active...... 15

3.1 Average roll rate comparison in regulation setup for the aileron case...... 18 3.2 Average roll rate comparison in landing configuration for the aileron case...... 19 3.3 Average roll rate comparison in cruise configuration for the aileron case...... 20 3.4 Average roll rate comparison in regulation setup for the spoilers case...... 23 3.5 Average roll rate comparison in landing configuration for the spoilers case...... 24 3.6 Average roll rate comparison in cruise configuration for the spoilers case...... 25 3.7 Average roll rate comparison in regulation setup with addition of rudder...... 26 3.8 Average roll rate comparison in regulation setup with enlarged ailerons...... 28

4.1 Actuator types and hydraulic line specifications for each control group in the weight model...... 32 4.2 Model component mass overview...... 34 4.3 Hydraulics system weight after modifying roll control hydraulics...... 34 4.4 Top-level specifications of the Boeing 747-200...... 36 4.5 Initiator results for an aircraft with Boeing 747-200 specifications...... 36 4.6 Interpolated Initiator mass results for the Boeing 747-200 with savings percentages... 39

5.1 Roll control on common jet airliners...... 42 5.2 Top-level speciﬁcations of the Boeing 767-200...... 43 5.3 Initiator results for an aircraft with Boeing 767-200 speciﬁcations...... 43 5.4 Interpolated Initiator mass results for the Boeing 767-200 with savings percentages... 45

NOMENCLATURE

CDR Roll drag coefﬁcient [-]

Clr Rolling moment due to yaw rate coefﬁcient [-]

Clβ Rolling moment due to sideslip coefﬁcient [-] L A Rolling moment around X-axis [Nm] m Aircraft mass [kg] Pss Steady-state roll rate [rad/s] 2 Sht Horizontal tail surface area [m ] 2 Sw Wing surface area [m ] 2 Svt Vertical tail surface area [m ] VTAS True airspeed [m/s] xcg Longitudinal location of the center of gravity [% MAC] yD Roll drag moment arm [m] φ Aircraft bank angle [degrees]

xiii

1 INTRODUCTION

This starting chapter aims to provide a solid foundation for the report. First a general context of the project is presented. In the second section the thesis goals and research question are stated. In the third section the basics of the ﬂight control system of the Boeing 747 are explained, together with a redundancy analysis. Lastly the structure of the report is given.

1.1. CONTEXT In commercial aviation there is a constant search for both increasing profit and decreasing environ- mental impact. The key parameter that affects this combination is fuel efficiency. New technologies that strive to improve fuel consumption on large airliners are always embraced by the industry and often spark a chain of improvements. Saving fuel can take many shapes and forms: better engines, better aerodynamics and lighter aircraft are just some of the approaches. The latter, decreasing the weight of the airframe, usually does not happen in big steps. It is more a process of new materials and shaving off a few components here and there. The flight control system is one of the subsystems where it may be possible to leave some parts out. This is not easy, because the safety and handling of the aircraft are at stake. Flight control mechanisms are in general complex and heavy because of safety through redundancy. There are usually many control surfaces present with identical functions e.g. rolling the aircraft. With this in mind there can be opportunities to leave out some superfluous systems to improve the aircraft weight and thus its fuel consumption. It is important however that safety is not compromised. Therefore only control surfaces with multiple backups should be considered. Other safety mechanisms like intelligent flight control systems can also be used, but this is outside the scope of this research project.

1.2. RESEARCHQUESTIONANDTHESISGOALS The main research question of the project is formulated as follows:

What unconventional ﬂight control possibilities are feasible on civil aircraft and what is their impact on ﬂight performance and weight?

The research question is the starting point for a collection of thesis goals. They are formulated below.

• To explore unconventional ﬂight control on large airliners

• To ﬁnd what unconventional ﬂight control methods are possible on the Boeing 747-200

• To assess which control surfaces can be left out while the aircraft is still able to ﬂy adequately

1 • To evaluate the limits and beneﬁts of the new ﬂight control laws through simulation and calculation

• To assess what parts can be left out of an aircraft ﬂying with the new control laws

• To estimate the direct weight impact of eliminating said parts

• To estimate the weight reduction on a newly designed aircraft, similar to the reference aircraft, but with the new conﬁgurations

• To extrapolate the results to other aircraft in the airline industry

1.3. BOEING 747 FLIGHT CONTROL SYSTEM The reference aircraft in this report is the Boeing 747-200. The reason for this specific type of aircraft is two-fold. Firstly it is not too recent and it was very popular in its time. As a consequence there is an abundance of technical knowledge and literature available for the public. Secondly, it is the subject aircraft of the RECOVER Benchmark, the software package that is used for the flight performance analysis. Details about this tool are explained later in the report. It is important to note that the flight control systems of this specific generation of the 747 family are not identical to the current model’s design (747-8). Changes include for example fly-by-wire outboard ailerons and aileron droop in cooperation with flap system. Even though the study in this report may seem outdated for this reason, it can be seen as starting point for further research. A thorough analysis on a more recent aircraft is nearly impossible without confidential information from manufacturers. The flight control system of the Boeing 747-200 is one of the most complex on any airliner. There are multiple layers of physical redundancy through multiple surfaces with identical functions and four different hydraulic systems powering various sets of control groups [1]. Being one of the most redundantly actuated aircraft in the airline industry, it is considered to be a good candidate to improve the structural weight at the cost of some redundancy.

Figure 1.1: Control surface layout on a Boeing 747. [1]

2 Figure 1.1 depicts an overview of the flight control surfaces on the Boeing 747. An extraordinary feature is the split rudder. To this day the 747 is the only airliner with this trait. Other aircraft designs have sufficient redundancy on this surface solely by multiple actuators, but have only one physical rudder. Nonetheless, compared to other control subsystems on the Boeing 747, the rudder group is still the most likely to fail (see Table 1.1). This is logical when considering the amount of control surfaces around other control axes. The excessive redundancy on the Boeing 747 makes it an extremely reliable aircraft, as has been proven numerically and practically. In legal terms, a flight control sys- 9 tem on a civil airliner must have a failure probability of less than 10− per flight. As can be seen in Table 1.1, the PFCS (primary flight control system) has a failure probability of nearly one tenth of this measure. Of that probability the rudder — which is in itself not critical to stay airborne — takes up 92.2%. In practice there are no records of major incidents caused by a failing flight control system (except due to considerable external damage) in the 747’s entire 47 years of service.

Table 1.1: Failure probability of individual control systems for a Boeing 747. [1]

Control system Probability of failure during 4-hour ﬂight 10 Elevator 0.218 10− ∗ 10 Lateral (aileron and spoiler) 0.211 10− ∗ 9 Rudder 0.107 10− ∗ 9 PFCS (elevator, aileron and rudder) 0.132 10− ∗

1.4. REPORTSTRUCTURE The report is organized as follows: Chapter 2 starts the exploration of unconventional flight control on civil airliners. Chapter 3 focuses on the flight performance impact of modifying the roll control systems. In chapter 4 the weight impact of the modified flight control system is assessed. Chapter 5 presents the application to other airliners. Chapters 6 and 7 contain the conclusions and recommendations, respectively.

2 UNCONVENTIONAL FLIGHT CONTROL

This chapter explores the unusual possibilities of ﬂight control surfaces on civil airliners. The main goal is to ﬁnd opportunities for further investigation.

2.1. UNCONVENTIONALCONTROLPOSSIBILITIES This section aims to give an overview of control surfaces and other mechanisms that could potentially be used as alternatives to the conventional surfaces. An overview of the discussed control possibilities can be found in Table 2.1. They are discussed in detail in the following subsections.

Table 2.1: Unconventional control possibilities around aircraft axes.

Axis of rotation Conventional control Unconventional control possibilities Roll (X) Ailerons (in-board/out-board), spoilers remaining ailerons, spoilers, differential elevator, flaps/slats Pitch (Y) Elevator Thrust, ailerons, flaps/slats, elevator trim Yaw (Z) Rudder Differential thrust, flaps/slats, ailerons, spoilers

2.1.1. ROLLCONTROL The principle of rolling an aircraft is based on creating a difference in lift force between its left and right side. Conventionally, this is done with ailerons: trailing edge tabs, usually as far from the fuselage as possible. Most aircraft have one set of ailerons, with the exception of some large jet airliners (e.g. Boeing 747, 767, 777) which also have an extra pair located more inboard. These inboard ailerons are used in high-speed roll maneuvers to alleviate wing loads and prevent aileron reversal. In large airliners like the Boeing 747 spoilers are also used for rolling. These surfaces (also called lift dumpers) disturb the airflow over the wing and cancel lift generation. It is an effective way to ’drop’ one of the wings, inducing a roll. Several alternative options exist for rolling an aircraft without the conventional ailerons. The first option would be to use the remaining ailerons, in case only one of multiple (inboard/outboard) sets is removed. On smaller aircraft where it is not part of the conventional roll mechanism, the next option would be to use the spoilers. These surfaces are located at a substantial distance from the aircraft’s X-axis thus it is easy to create large moments. The downside of using this method would be the increase in drag. The aircraft will lose airspeed and will undergo a yawing moment. The third option, differential elevator, is more revolutionary than the previous two. The idea is to deflect the elevators to opposite sides, like ailerons. Since this is still a small trailing-edge deflection it is quite

5 efficient in terms of drag. However, the torque obtained from this method is expected to be small because of the short moment arm around the X-axis. The last option would be use high-lift devices. This is the complete opposite concept of spoilers. By deploying flaps and/or slats on one wing, that side is lifted and the aircraft rolls. The major issue with this method is deployment speed of the surfaces. In current aircraft, deployment speed is of no importance so the actuators are small and weak. This method’s feasibility highly depends on the type of high-lift devices present. For example, Krueger flaps like on the Boeing 747 would be useless. Lightly rolling would be impossible because of the binary states of the devices — they are either deployed or not, without an intermediate state.

2.1.2. PITCHCONTROL Rotation around the Y-axis is conventionally done with elevators. These trailing-edge tabs on the horizontal stabilizer are positioned as far from the center of gravity as possible. In some rare cases canard surfaces are used, sometimes as a substitute for elevators and sometimes in combination (three-surface aircraft). However,t his concept is rarely present in airline aviation so it will not be discussed further. Controlling pitch of the aircraft without elevators is not easy as there are not many movable surfaces at a significant distance from the Y-axis. One way to gain some control over the aircraft is with the engines. This method has been studied and proven before [3]. Assuming the non-moving part of the horizontal stabilizer is still intact, a pilot or flight control computer can use the aircraft’s longitudinal stability to control pitch. It is rather straightforward: increase in speed induces a nose up moment, and vice versa. The major downside of this method is the response delay. When a thrust increase command is given, the engines have to spool up which can take up to several seconds. After that there is even more time necessary for the aircraft to increase angle of attack and lift its nose. At altitude this of little concern but when landing this would be very undesirable. The second option, using ailerons, should provide a better response time, albeit at the cost of effectiveness. With swept wings the ailerons are located at a certain distance behind the center of gravity. Deflecting both ailerons in the same direction will create a moment around the Y-axis. Depending on wing parameters and aileron size this may offer a feasible solution. Thirdly, the aircraft’s high-lift devices may offer (limited) solutions. As mentioned in the previous subsection, these surfaces move very slowly. Additionally, they can only provide a nose-down moment. This makes it essentially a solution for the sole case of the elevators being stuck in a nose-up position. Lastly, the elevator trim mechanism may be of use. On large airliners this mechanism usually consists of an all-moving tail. It has limited and slow angular movement but since it moves the entire tail surface, it can have a decent influence on the aircraft attitude.

2.1.3. YAWCONTROL The third and last control axis is the Z-axis. Control around this axis is normally achieved with a rudder. This is again a trailing edge tab, this time on the vertical stabilizer. Deflecting this surface essentially pushes the tail to one side. Aircraft with tall vertical tails also experience a certain roll moment due to this deflection but the dihedral angle of the wings can balance this perfectly. The yaw axis is considered to be the least critical axis. An aircraft is still flyable without rudder even though making a turn will be less efficient. The vertical tail of modern airliners is designed to keep the aircraft stable without rudder, even with one side’s engines off. There are no other surfaces in the same plane as the rudder, so for unconventional control other options need to be explored. The main concept is to create more drag on one side of the aircraft than on the other. This can be done with several surfaces at once, even with a combination of differential engine power. The difficulty of this process lies in the fact that all these extra distortions need to cancel out around the other axes. For example, if the goal is to yaw rightward, the control system could deploy spoilers on the right wing. To cancel out the loss in lift on this wing the high-lift devices and or

6 aileron could be deployed (which do their part as well for creating extra drag). This is a complex and slow balancing act without use of complex ﬂight computers and ﬂy-by-wir. It is nearly impossible as effective control method, especially when executing critical maneuvers near ground level. It could however be used to trim the aircraft in case of engine and rudder failure.

2.2. ANALYSISOFNEWCONTROLOPPORTUNITIES From the previous section the most promising control methods are chosen and analyzed in below. The selection is limited since some of them were disregarded because of limitations in the simulation tool.

2.2.1. RECOVERBENCHMARK The performance analysis tool in this project is RECOVER Benchmark (REconfigurable COntrol for Vehicle Emergency Return) by GARTEUR (Group for Aeronautical Research and Technology in EU- Rope). It is a software package in the Matlab/Simulink environment (see Figure 2.1), specifically designed for reconfigurable flight control. The full extent of the project that worked towards this tool is explained in [4]. In short, it is an extensive, non-linear simulation program for the handling performance of a Boeing 747-200 freighter. It offers control to everything a real-life pilot would have access to, ranging from engine power to flaps and trim tabs. It can also be equipped with advanced control laws, including intelligent reconfiguration algorithms. The simulated aircraft can then be put through any type of scenario, from cruise flight to disaster cases. Since its release, it has been success- fully used in many research projects (e.g. [5]), proving its value. The main basis and key motivation for the choice of this particular aircraft is the disaster of El Al Flight 1862. In 1992 a 747-200 freighter crashed near Amsterdam after its two right-wing engines were separated from the wing. Several re- ports, including the official crash report, have stated that the crash could possibly have been averted had there been a form of fault-tolerant flight control present on the aircraft. That theory has been proven to be plausible with this tool. [6–8] Currently the software package is mainly used for simulating intelligent control laws. Its purpose for this project on alternative flight control is primarily an aerodynamic analysis tool. It was chosen over other aerodynamic software packages because of completeness and ease of access. Simulating the resulting forces and moments of a single control surface deflection would technically be possible with CFD software or other aerodynamic simulation tools, but evaluating multiple control surfaces in specific conditions on an entire aircraft would be impossible to do in the time frame of this project. Additionally, the subject of the Benchmark tool, the Boeing 747-200, is one of the most redundantly actuated aircraft in the airline industry which is perfect for the exploration of new control methods. It should be noted that the aerodynamics in the Benchmark tool are limited. Although the overall complexity is rather high because of the amount of variables and the size of the aircraft, all calculations are based on graphs such as a CLα graph. As a consequence, complex aerodynamic effects in for example the high trans-sonic region are not modeled perfectly and the results should always be received with some criticism.

OPERATING METHOD Except for the initial boot screen of the Benchmark tool, it supports no graphical user interface. In that initial screen, there are two main options to choose from: open-loop simulation and closed-loop simulation. For this project only open-loop is used since no auto-pilot or other controllers are used. All interactions with the software after selecting this option are done through Matlab prompt in the command window. The software asks step by step what the user needs to specify for the simulation, as shown in Figure 2.2. First the trimming sequence runs. In this part the software prompts properties of the aircraft (mass, operative engines,...) and of the current maneuver (steady climb, level turn,...). With this information the software will then optimize the aircraft’s trim. This was found to

7 8

Figure 2.1: General Simulink structure of the RECOVER Benchmark. work adequately, with the exception of low-speed ﬂight. Even thought the ofﬁcial stall speed of the aircraft is around 60 m/s, the trimming algorithm is not stable below 75 m/s.

Figure 2.2: Typical command window prompt of RECOVER Benchmark

After the aircraft’s trim state is found, the interface prompts what the pilot’s input should be for the following, user-specified time. This input can be any combination of the standard pilot controls, defined by an input function e.g. step or sinus function. This is the final step before the program puts out an extensive amount of graphs, covering all states and variables from start to finish. However, this step is not going to be used in the project, as it is limited to only conventional pilot inputs. Deflecting different control surfaces individually is not possible through this method so it is bypassed. To bypass the regular operation of the Benchmark tool, some changes were made to the Simulink model. After careful analysis of the inner workings of the model, it was decided that the most ro- bust way controlling individual surfaces is to simply interrupt the signals deep inside the "autopilot" block and replace them with input functions. An example of this procedure on the elevator controls can be seen in Figure 2.3. Note that these ’blocks’ are buried five layers deep into the model so the modifications were rigorously tested to make sure that all possible effects were accounted for.

Figure 2.3: Modiﬁcations in RECOVER Benchmark to control the elevator surfaces individually

2.2.2. PITCHCONTROL As mentioned in subsection 2.1.2 the options for pitch control besides elevators or canards are limited. Due to limitations of the aircraft model in RECOVER Benchmark and the fact that it has been used in practice before, thrust variation will not be considered in this report. That leaves one viable option left: ailerons. Several setups and scenarios can be considered to validate these control surfaces. From a regulations standpoint there are no numerical requirements for an aircraft pitch command. For other maneuvers like roll and yaw there are certain numerical constraints described

9 in EASA CS-25 documents, as will be discussed in the next subsection. It is therefore imperative to set up a decent framework in order for different cases to evaluated. Within the image of a standard airline flight one critical pitch maneuver comes to mind: the landing flare. It would be only logical that during this specific phase the force required by the pitch- regulating surfaces is the highest. Several factors contribute to this, the first being the low airspeed. The control surfaces should be large enough and have enough deflection potential in order to generate enough aerodynamic force at approach velocity. Secondly, the full deployment of high-lift devices has a negative influence on the intended maneuver. Flaps and slats (regardless of type) generate a substantial nose-down moment. Besides requiring the heaviest load during flight, this maneuver is also the most critical in terms of safety. If anything would go wrong during the flight, the only thing that is absolutely critical is getting the aircraft safe on the ground. The less the aircraft can perform a landing flare, the more it becomes a crash instead of a landing. To simulate the specific case of a landing flare, the aircraft model is trimmed in landing state. This involves a flight path angle of -3 °, approach speed of 77 m/s, maximum landing weight and the most stable center of gravity. The parameters are the following:

• m 285763 kg =

• x 0.11 MAC cg =

• Flaps deployed max (30)

• Landing gear deployed

• Altitude = 600 m

• V 77 m/s TAS =

• γ 3o = −

Running the simulation with aileron inputs instead of elevators provided remarkable results. First off, the influence of the inboard ailerons was found to be negligible. This was to be expected because of the placement and size of the surfaces. The longitudinal distance from the center of gravity (even in its most forward position) to the inboard ailerons is little more than three meters. Combining this with the small surface area and deflection limits, leads to a very small moment. After all, these surfaces were designed for light maneuvers at high speed. Consequently, at landing approach speed they fall short for both roll and pitch control. Contrary to their inboard equivalents, the outboard ailerons show some promising results. The test simulation consists of two parts. The first five seconds the aircraft is flown in its trim condition. This is to smooth out irregularities from the software’s automatic trim function. Then, after five seconds, the given input command is executed. After 20 seconds the simulation is terminated. Since the input is always a simple step function, any results after a long time become increasingly irrelevant because it would never occur in reality.

10 Figure 2.4: Altitude and Alpha plots with outboard aileron input during ﬁnal approach: full upward deﬂection

As can be seen in Figure 2.4 it is possible to lift the nose of the aircraft up in a reasonable amount of time. It should be stressed that this is with full upward deflection of the outboard ailerons. This would mean that they are not usable for roll maneuvers during this phase. However, the Boeing 747 makes use of spoiler deployment as standard while landing for roll maneuvers and they suffice on their own, as shown later in the report. For comparing purposes, Figure 2.5 shows what a normal pilot input would result in and Figure 2.6 shows what the aircraft’s response would be if it stayed in the initial trim state. It should be noted that a so-called ’normal input’ is around 50% of the maximum limit on the control column. This was derived from pilot experiences. From these graphs it is clear that in normal operation the aircraft levels off in about 5 seconds after elevator deflection. In the 8 seconds thereafter a steady climb occurs, which stagnates in the final seconds of the simulation. In the last few seconds the aircraft starts entering stall, due to the increasing angle of attack and dropping airspeed. Note that the engine settings are kept in their initial (descent) trim state so it is sensible that the airspeed decreases significantly when climbing again.

Figure 2.5: Altitude and Alpha plots with normal pilot input during ﬁnal approach: -6 degrees on control column (50%)

11 Figure 2.6: Altitude and Alpha plots without any control input during ﬁnal approach

Although it is surprising that ailerons can have such an influence on pitch attitude, there is some sensible reasoning to it. These causes are specific to the Boeing 747 models, making it unlikely that such favorable results could be replicated on other airliners. Firstly, the outboard ailerons are quite large. At an area of 1.5 m2 each, they are almost double the size of the inboard ailerons and take up a substantial amount of the wingspan. Secondly, the sweep angle on the Boeing 747 is unusually large for this type of aircraft. As Table 2.2 shows, the only other common airliner that comes close to this property is the Airbus A350. All other aircraft have significantly lower sweep angles.

Table 2.2: Wing quarter-chord sweep angles for common jet airliners

Aircraft Wing sweep angle [°] Boeing 737 25 Boeing 747 37.5 Boeing 757 25 Boeing 767 31.5 Boeing 777 31.64 Boeing 787 32.2 Airbus A320 25 Airbus A340 30 / 31.1 Airbus A350 35 Airbus A380 33.5

The above results on alternative pitch control are surprising, but nonetheless they are not sufficient enough to initiate further exploration. With full deflection the benchmark is only just met, which means that more extreme circumstances are not feasible. The hypothetical next step would be to eliminate the elevator surfaces. Safety and handling margins would demand for a better response and that is not feasible with the outboard ailerons, without any major modifications. They are already places as far back as possible and enlargement (as will be done later in the report for roll control) is not a profitable approach. Stretching ailerons span-wise would have marginal return because the aerodynamic force would shift forward. Stretching chord-wise, although more profitable, is limited by wing structures and thus offers limited opportunity. Therefore the investigation on alternative pitch control is not continued after this point.

2.2.3. ROLLCONTROL In contrast to pitch control, roll maneuvers offer many alternative control methods as discussed in subsection 2.1.1. The possibilities include ailerons, high-lift devices, spoilers and elevators. However,

12 due to limitations of the RECOVER Benchmark software, elevators and high-lift devices will not be evaluated in this report. Unfortunately the aerodynamic model in the tool is incapable of calculating inﬂuences of single-sided deployment of these surfaces.

REGULATIONS When looking into flight control methods, it is important to know what the governing bodies dictate as limits and requirements. As mentioned before, for pitch control there are no numerical or exactly defined limits. Yet, rolling is a different story. It is convenient to define an aircraft’s roll rate and as such, the European Aviation Safety Agency (EASA) has set a certain minimum for each type of aircraft. The exact conditions under which this requirement is tried and tested, are defined rigorously. The literal statement taken from CS-25 [9] can be found below.

EASA CS-25.147(d): With the aeroplane in trim, all as nearly as possible,in trim, for straight ﬂight at V2, establish a steady 30° banked turn. It should be demonstrated that the aeroplane can be rolled to a 30° bank angle in the other direction in not more than 11 seconds. In this demonstration, the rudder may be used to the extent necessary to minimize sideslip. The demonstration should be made in the most adverse direction. The manoeuvre may be unchecked. Care should be taken to prevent excessive sideslip and bank angle during the recovery. Conditions:

• Maximum take-off weight.

• Most aft c.g. position.

• Wing-ﬂaps in the most critical take-off position.

• Landing Gear retracted.

• Yaw SAS on, and off, if applicable.

• Operating engine(s) at maximum take-off power.

• The inoperative engine that would be most critical for controllability, with the propeller (if applicable) feathered.

With the above conditions EASA believes that the most critical phase in ﬂight for a roll maneuver is take-off. With the aircraft trimmed accordingly it must be able to roll from a 30 degree bank angle on one side to 30 degrees on the other side within 11 seconds. Testing this scenario in RECOVER Benchmark follows the same procedure as before with pitch maneuvers. After trimming the aircraft with the parameters below, the input is speciﬁed.

• m 351535 kg (MTOM) = • x 0.33 MAC cg = • Flaps deployed T/O (20/30)

• Landing gear retracted

• Altitude = 600 m

• V 93.5 m/s TAS = • φ 0.5812 rad (30 degrees) = 13 For reference purposes the normal, conventional setup is evaluated. This involves a step input to the control wheel at maximum deﬂection (88°/1.53 rad). The roll angle response of the virtual Boeing 747-200 can be found below. At t 12 (11 seconds after the step input starts) the aircraft has rolled = 44 degrees to the left of its straight-up position. This means it has met the 30 degree requirement by EASA, but not by much. The time stamp at which it surpasses the 30 degree mark is only 1.19 seconds short of the limit. There is one thing to note that acts as an attenuating circumstance and that is the rudder trim. The start position of the aircraft is trimmed for a continuous turn at 30 degrees bank angle. The automatic trim function of the software uses the rudder to keep the aircraft in the most optimal position. In this case, the rudder is deﬂected for one third of its mechanical limit and this has a substantial impact on the aircraft’s attitude.

Figure 2.7: Result of a max input roll maneuver following EASA standards

SIGNIFICANCE OF SPOILERS When looking into roll maneuvers of the Boeing 747-200, it is striking how many control surfaces are used for simple actions. In regular flight a combination of inboard ailerons and spoilers is used. Not all spoilers are active; the two most inboard are only used for air braking and lift dumping. This is likely a consequence of the trade-off between the additional drag and the little rolling moment these surfaces can generate at the short distance from the aircraft’s center line. To be able to roll adequately at low speeds, the outboard ailerons are used as well. As soon as the flaps are deployed (starting from 1 unit out of 30) the outboard ailerons are activated. All these different setups raise the question how effective each of these surfaces is. The responses (expressed in bank angle φ) are plotted in Figure 2.8. In Table 2.3 the maximum achievable roll rates for different surfaces are quantified. These were tested in the setup defined below. The roll rate was measured over five seconds, starting 5 seconds after the input command to make sure most acceleration is out of the picture.

• m 317000 kg =

• x 0.25 MAC cg =

• Flaps retracted / deployed at 1/30 for last scenario

• Landing gear retracted

• Altitude = 600 m

• V 133.8 m/s TAS =

• Straight and level ﬂight

14 Figure 2.8: Roll response comparison between scenarios with different surfaces at maximum deﬂection

Table 2.3: Roll rate comparison between scenarios with different surfaces active

Active Surfaces Average Roll Rate [deg/s] Inboard ailerons 3.3 Inboard ailerons + spoilers 8.7 Inboard ailerons + spoilers + outboard ailerons 15.3

The most important thing to note is how little the inboard ailerons contribute to the total roll rate in normal ﬂight. The spoilers take up nearly 60% of the workload during a roll maneuver. This remarkable division of contribution sparked an interest in possible innovations for large airliner wings. Maybe it is possible to design an aircraft with less control surfaces and to save weight or improve other features? The elaborate discussion on the up- and downsides of these choices continues in the next chapters.

3 FLIGHTPERFORMANCEIMPACTOFFEWER CONTROLSURFACES

The exploration of the different control mechanisms of the virtual Boeing 747-200 of RECOVER Bench- mark triggered questions about the efficiency and necessity of some control surfaces. Especially roll maneuvers may contain opportunities for improved designs. This chapter presents two cases concerning the change in flight performance if the aircraft would be designed with a different control surface layout. The first case is about the reduction of ailerons; either only inboard ailerons or all ailerons. The second examines the other side: leaving out spoilers. The two cases are evaluated in various circumstances: regulation-specified, landing and cruise.

3.1. CASE 1: FEWERAILERONS The ﬁrst case is perhaps the most revolutionary. There are aircraft with all kinds of extraordinary designs, but aircraft without ailerons are rare. The idea that led to this possibility was the small contribution inboard ailerons have on the roll rate of the Boeing 747-200, found earlier in the project.

3.1.1. REGULATIONS The ﬁrst step in evaluating new control setups is checking with regulations. There would be no use for further exploration if regulation standards cannot be met. The roll regulations, from which the simulated conditions are derived, were explained in subsection 2.1.1. To keep things simple and clear from now on the maneuvers will be compared by roll rate only. The requirement is rolling from 30 degrees on one side to 30 on the other within 11 seconds. This translates to 60 degrees in 11 seconds. The average roll rate is then 5.5 degrees per second, which will be the number to beat. For clarity the aircraft’s trim for regulation testing is repeated below. The evaluation results of the Boeing 747-200 with various aileron setups are summed up in Table 3.1.

• m 317000 kg = • x 0.25 MAC cg = • Flaps retracted (0/30)

• Landing gear retracted

• Altitude = 11000 m

• V 93.5 m/s = 17 Table 3.1: Average roll rate comparison in regulation setup for the aileron case

Conﬁguration Average roll rate [deg/s] Normal operation (all ailerons + spoilers) 11.7 No inboard ailerons 7.5 No outboard ailerons 6.6 Only spoilers 4.8

From Table 3.1 it is clear that even at low speeds there is a definite impact from the inboard ailerons. Nonetheless, with these surfaces out of the picture regulation standards can been met easily. To continue the trend of reducing the amount of control surfaces located on the wing, the lack of outboard ailerons is also evaluated. Theoretically this decision should save a significant amount of weight, which will be proven later on in REFERENCEEE. As expected, their influence at low velocities is greater than the inboard ailerons. However, the difference is only around 1 degree per second. The aircraft would still be legal to fly without outboard ailerons. The extreme case of no ailerons at all poses a problem: a roll rate of 4.8 degrees per second is 0.7 units short of the lower limit imposed by EASA (5.5). Although the total surface area and the deflection limit of the spoilers is much greater than that of the ailerons, the roll rate drops 60 % when all ailerons are eliminated. The explanation lies in the fact that only half of the spoilers contribute to the workload at once. When rolling to the left, only the left wing’s spoilers are used while the right-hand wing remains unchanged. Consequentially there is only a force acting downward on one side and no change in lift on the other which is not the most efficient way of rolling an aircraft. A possible solution is adding more spoiler surfaces. This issue is addressed in subsection 3.1.4. Additionally, compared to ailerons the spoilers are more airflow-disturbing. The drag generated from this is undesirable in long maneuvers. This is explained more in subsection 3.1.3.

3.1.2. LANDING One of the more critical phases in flight is without doubt the landing phase. During this maneuver it is important to have enough roll capabilities to counter for example side wind gusts. Close to the ground there is no time or room for the aircraft to stabilize itself. To test the capabilities of the aircraft the model was trimmed in the configuration below. The findings are displayed in Figure 3.1 and Table 3.2.

• m 255800 kg (MLW) =

• x 0.25 MAC cg =

• Flaps deployed fully (30/30)

• Landing gear deployed

• Altitude = 600 m

• V 85 m/s TAS = 18 Figure 3.1: Maximum achievable roll maneuver comparison in landing setup for the aileron case

Table 3.2: Average roll rate comparison in landing conﬁguration for the aileron case

Conﬁguration Average roll rate [deg/s] Normal operation (all ailerons + spoilers) 7.7 No inboard ailerons 5.4 No outboard ailerons 5.0 Only spoilers 3.7

As anticipated, the overall roll performance has decreased when compared to the regulation con- ﬁguration (see Table 3.1). Due to the lower airspeed the maximum achievable roll rate is now only 7.7 degrees per second. The impact of leaving out either inboard or outboard ailerons is similar in this case, dropping the roll rate by a quarter approximately. Eliminating both sets of ailerons at once has a more dramatic impact of about 50%. As stated before, at low velocities near ground level there is not much room to spare concerning maneuvering capabilities. Therefore this performance is deemed unsatisfactory.

3.1.3. CRUISE The last conﬁguration to be considered is when the aircraft is in cruise phase. Although cruise is generally in a straight line, direction changes are not uncommon. It also gives a general indication of the performance during airport approach. The trim setup was chosen to represent normal mid-cruise conditions as close as possible. They are listed below. The difference in roll response is depicted in Figure 3.2 and numerical roll rate results can be seen in Table 3.3.

• m 317000 kg =

• x 0.25M AC cg =

• Flaps retracted (0/30)

• Landing gear retracted

• Altitude = 11000 m

• V 263 m/s (M 0.89) TAS = = 19 Figure 3.2: Maximum achievable roll maneuver comparison in cruise setup for the aileron case

Table 3.3: Average roll rate comparison in cruise conﬁguration for the aileron case

Conﬁguration Average roll rate [deg/s] Normal operation (inboard ailerons + spoilers) 19.1 Only spoilers 11.9

As shown above, only two different setups are considered. Because of aileron reversal and wing loading the outboard ailerons are not used in cruise anyway, which leaves only the option of inboard ailerons. When leaving the latter locked in place, the roll performance drops by 35%, which is a significantly larger impact than in the previous flight phases (comparing "No outboard ailerons" and "Only spoilers". 11.9 degrees of roll rate should still be more than enough during most — or even all — maneuvers. However, it should be noted that this maximum roll rate is achieved with full deflection of the spoilers (45 degrees). This amount of flow disturbance is not favorable and will likely cause unexpected flight behavior. The advanced aerodynamics involved may also be modeled poorly RECOVER Benchmark. The authenticity and accuracy of the results are therefore not guaranteed.

3.1.4. MODIFICATIONS

In subsection 3.1.1 it was shown that the legal roll requirement could not be met with just the spoilers. Without resorting to the ailerons, a solution to this issue is adding spoilers. Even though this approach adds weight, it is likely to still be lighter than keeping a set of ailerons. Adding identical spoiler surfaces next to the existing set is the most straightforward approach. To maximize effectiveness, they should be placed as far outboard as possible. Since the ailerons would be eliminated in this case there are no spacial limits in this direction. To assess how many extra spoilers would be necessary a rudimentary extrapolation can be made from the existing surfaces. It should be noted that not all spoiler surfaces are identical on the Boeing 747-200. As visible in Figure 3.3, the inner two ailerons are separated from the outer four. The inner pair (nr. 5 and 6) have a mechanical deflection limit of 20 degrees, whereas the outer four (nr. 1 through 4) can deploy up to 45 degrees. The flight control mechanism of the Boeing 747-200 does not use spoiler nr. 6 (and nr. 7 on the right side) for rolling. Because of the deflection limits and location, the data points from which the extrapolation can be made are only of spoilers nr. 1 through 4.

20 Figure 3.3: Wing layout of the Boeing 747-200.

The cumulative roll rate of the spoiler surfaces is depicted in Figure 3.4. Not that the numbers on the X-axis do not coincide with the spoiler identification numbers mentioned before. When two ailerons are being used, that means spoilers nr. 5 and 4 are deployed fully. Three ailerons in use means nr. 5, 4 and 3; and so forth. Following the trend line towards 6 spoilers, it is visible that the regulations are met, albeit barely. To have some assurance and safety margin, a total of 7 rolling spoilers would suffice, bringing the total to eight spoilers per wing. With this modification the average roll rate of the aircraft becomes nearly 6 degrees per second. Logically, adding spoiler surfaces has some weight penalty, which will be discussed later in the report.

Figure 3.4: Effectiveness of adding spoilers towards the wingtip.

3.1.5. ADDITIONALREMARKS An important consideration of using spoilers — especially at high speeds — is their efficiency. It has been mentioned previously that because of their location (on top of the airfoil) and their large deflection angle these surfaces have quite an impact on the airflow. In order to quantify this impact

21 a comparison between total energies has been produced (shown in Figure 3.5). The total energy in the graph is calculated through the sum of the potential energy and the kinetic energy. It is assumed that at a zero roll rate the aircraft’s energy stays constant, which makes it convenient for a fractional comparison. From the graph it becomes clear that rolling with ailerons only causes a loss of less than 1 % of the aircraft’s energy after 10 seconds, at roll rates up to 7 degrees per second. On the contrary, rolling with spoilers at the same rate makes the aircraft lose nearly 3.5% of it’s energy over just 10 seconds. When cruising this would result in a loss of 5 m/s in that time frame. Although this is a significant result, it is not surprising. After all, the spoilers are also used as speed brakes. This efficiency concern, together with the unsatisfactory performance during the landing phase make the option of using only spoilers on the Boeing 727-200 invalid. With some modifications however, this is still an interesting possibility. Extending the spoiler surface outboard would yield increasing rolling moments due to the further distance from the X-axis.

Figure 3.5: Efﬁciency comparison between aileron roll and spoiler roll maneuvers

The second consideration that comes up when flying with only spoilers, is safety. As mentioned before, the aircraft does not fly like a conventional aircraft in the way that it just drops one wing. As a consequence, the aircraft will not be able to fly a turn as close to the ground and as slow as a normal aircraft would. This is important for pilots to know — or for the flight envelope protection on fly-by-wire aircraft — when performing more critical maneuvers near take-off and landing.

3.2. CASE 2: NOSPOILERS The second case explores the opposite opportunity of case 1. Instead of leaving out the ailerons, the performance impact of ﬂying without spoilers is evaluated.

3.2.1. REGULATIONS As was the case with no ailerons, the ﬁrst evaluation of the aircraft without spoilers is within the regulatory frame. The requirements are obviously the same as before, since the subject is still a roll maneuver. In simpliﬁed terms it comes down to having a minimum roll rate of 5.5 degrees per second. The aircraft trim is repeated below and the results can be found in Table 3.4. The simulated results are displayed below, in Figure 3.6 and Table 3.4.

• m 351535 kg (MTOM) = • x 0.33 MAC cg = 22 • Flaps deployed T/O (20)

• Landing gear retracted

• Altitude = 600 m

• V 93.5 m/s TAS =

Figure 3.6: Maximum achievable roll maneuver comparison in regulation setup for the spoilers case

Table 3.4: Average roll rate comparison in regulation setup for the spoilers case

Conﬁguration Average roll rate [deg/s] Normal operation (all ailerons + spoilers) 11.7 All ailerons 4.4

The ﬁrst thing to notice is that the roll rate with only ailerons is too low. 4.4 degrees per second is about 20% short of the legal requirement. This poses a serious problem for this case. On the other hand, it proves the purpose of the spoiler surfaces in roll maneuvers and shows why it is not uncommon in large airliners. In order for the aircraft to be airworthy it is mandatory to ﬁnd a solution to the low roll rate. Two possibilities are discussed in subsection 3.2.4.

3.2.2. LANDING For the landing phase the aircraft is once again trimmed like before. The settings are displayed below. The simulation results are shown in Figure 3.7 and Table 3.5.

• m 255800 kg (MLW) = • x 0.25 MAC cg = • Flaps deployed fully (30/30)

• Landing gear deployed

• Altitude = 600 m

• V 85 m/s TAS = 23 Figure 3.7: Maximum achievable roll maneuver comparison in landing setup for the spoilers case

Table 3.5: Average roll rate comparison in landing conﬁguration for the spoilers case

Conﬁguration Average roll rate [deg/s] Normal operation (all ailerons + spoilers) 7.7 All ailerons 3.1

In the landing phase it is clear that there would be serious performance loss without spoilers. Less than half of the maximum achievable roll rate remains. This is not comfortable in a situation where quick responses are often required to deal with side gusts or general crosswinds.

3.2.3. CRUISE The trim properties of the aircraft in the cruise phase are repeated below. The results are displayed in Figure 3.8 and Table 3.6.

• m 317000 kg = • x 0.25 MAC cg = • Flaps retracted (0/30)

• Landing gear retracted

• Altitude = 11000 m

• V 263 m/s (M=0.89) TAS =

Figure 3.8: Maximum achievable roll maneuver comparison in cruise setup for the spoilers case

24 Table 3.6: Average roll rate comparison in cruise conﬁguration for the spoilers case

Conﬁguration Average roll rate [deg/s] Normal operation (inboard ailerons + spoilers) 19.1 Only inboard ailerons 9.0

The trend of losing more than half of the rolling performance when flying without spoilers continues into the cruise phase. However, a roll rate of 9 degrees per second is still more than enough in this stage of flight. Structural loads and passenger comfort limit the roll rate generally to about 5 degrees per second during most airline flights. It is not without risk to fly with a limited aircraft though. The reserves in performance and handling are generally present to compensate for failures and make the aircraft fail-safe.

3.2.4. MODIFICATIONS

Previously it was shown that without spoiler surfaces the roll performance of the Boeing 747-200 is lacking. Two possible solutions to this issue are considered, the first being the use of rudder to roll. This practice is common in modern fighter jets with fly-by-wire technology. These aircraft generally have the property that a certain pilot control deflection concurs with the same flight response, regardless of flight conditions. For example, if the pilot deflects the stick 5 degrees to one side, the aircraft’s flight computers will calculate the most efficient and convenient way to make the aircraft roll with X degrees per second. This roll output will always be striven for, whether the aircraft is flying slowly at take-off or fast in cruise.

The Boeing 747-200 — as well as other four-engine airliners — has a very tall and large vertical tail of which the rudder surfaces take up a considerable portion. At a height of 10 m and a surface area of around 20 m2 it can provide substantial influence to the aircraft’s attitude. However, this is not as straightforward as it may seem initially. The results of adding full rudder deflection to the concept of only ailerons can be seen in Figure 3.9 and Table 3.7. From the numbers only, it seems like the roll rate without spoilers is no longer an issue as it is larger than the required 5.5 degrees per second. It even surpasses the normal roll rate with spoilers by 1.5 units. However, this is obtained through indirect control instead of direct. The extra roll moment is a consequence of the two stability derivatives Clβ and Clr . The rudder is deflected negatively (to the right), pushing the nose of the aircraft to the right. Coupled with the aircraft’s dihedral angle, this motion creates a rolling moment to the right which adds to the aileron deflections. This all seems very promising but care should be taken with these results. Large deflections of the rudders will have a much greater influence on their intended purpose (directional control) than their secondary purpose. This can be quantified in terms of sideslip angle β, as shown in Figure 3.10. In phases like landing this may have catastrophic consequences. The maximum deflection angle of 25 degrees is designed for engine failures and other extreme cases that require large forces. In normal flight only values up to around 10 degrees are used.

25 Figure 3.9: Maximum achievable roll maneuver comparison in regulation setup with addition of rudder deﬂection

Table 3.7: Average roll rate comparison in regulation setup with addition of rudder

Conﬁguration Average roll rate [deg/s] Normal operation (all ailerons + spoilers) 11.7 All ailerons 4.4 All ailerons + rudder 13.2

Figure 3.10: Sideslip response from the aircraft’s rudder deﬂection

The second solution to the roll rate shortcoming is physically increasing the aileron effectiveness. This can be done in two ways: increasing the deflection angle and/or increasing the effective area. The former method may be the easiest to implement in many cases. In most aircraft a longer actuator arm would be sufficient. Despite being a convenient path to follow, it is not without risk. At extreme deflection angles in certain conditions the airflow will be disrupted severely with separation and buffeting as possible consequences. The second possibility is not prone to this drawback, but it requires more dramatic change. In the aircraft’s design phase, the sizing of ailerons is usually not the first priority. It comes second to the size of the high-lift devices and wing structure. In most aircraft the ailerons therefore neatly take up the left-over space between the wing’s edges, the rear spar and the trailing-edge flaps. Nonetheless, on the Boeing 747-200 there is an opportunity for this solution to work. Some quick analysis reveals there is still room to increase the size of the outboard ailerons. The inboard ailerons are closely boxed by the flaps and the wing box ’kink’. These elements are however not as constraining around the outboard ailerons. Span-wise they could be enlarged from 6.7 m to 8.5 m. This still leaves some space near the winglet for structural parts if necessary. Chord-wise an expansion from 0.9 m to 1.3 m is possible, right up to the rear spar. A graphical representation of these modifications can be seen in Figure 3.11 These modifications would increase the area of one outboard aileron from 5.9 m2 to 11.3 m2.

26 Figure 3.11: Expansion of the outboard aileron on a Boeing 747-200

s 2L A Pss (3.1) = ρ(S S Svt)C y3 w + ht + DR ∗ D

Without doubt, a control surface area increase of 92% has a significant impact. The theoretical roll rate method from Sadraey [10] has been used to prove this hypothesis. The equation stated above gives the relationship between the steady-state roll rate Pss and the forces involved. The numerator contains the total rolling moment around the X-axis L A and the denominator contains all factors contributing to the roll drag of the wing and tail. A few assumptions can be applied to this formula. The first is that the denominator does not change because of the aileron redesign. The second is that the lift force generated by a trailing-edge control surface is proportional to its surface area. This means that the increase of 92% in area results in a rolling moment increment of also 92%. From the formula below it then follows that the steady-state roll rate would increase by p1.92 1.39. Note that = this is only true for the outboard ailerons. The results of applying this knowledge to the regulations roll case without spoilers are displayed in Table 3.8. The contribution of the outboard ailerons is multiplied by 1.39 and the unmodified inboard aileron contribution is added to obtain a new roll rate of 5.5 degrees per second. This is a very interesting result because it reaches the EASA standard, albeit with zero margin. It should be noted that this is achieved only because the outboard ailerons were magnified up to what is thought to be the constructional limit. If any additional limits would be overlooked, this approach is not enough to make the aircraft legal to fly without spoiler surfaces. And even then, a barely legal product would not be attractive to the customer. It is deemed to be a plausible case, certainly with a combination of the aforementioned solutions: usage of the rudder and expanded outboard ailerons.

27 Table 3.8: Average roll rate comparison in regulation setup with enlarged ailerons

Conﬁguration Average roll rate [deg/s] All ailerons 4.4 Outboard ailerons 2.8 Enlarged outboard ailerons 3.9 Inboard + enlarged outboard ailerons 5.5

3.2.5. ADDITIONALREMARKS

In the previous sections it has been shown how the aircraft would be able to meet roll requirements. This is however not the only use of spoilers on a Boeing 747-200. In fact, the main reason why spoiler surfaces were ﬁrst placed on aircraft wings is for airbraking and lift dumping. Without spoiler surfaces these two functions are not available, which causes several issues. When descending from cruise altitude a jet airliner typically deploys its spoilers. This allows for steeper descent angles without exceeding the dive speed. This functionality is especially desirable for military aircraft which would form easy targets when ﬂying low. Jet airliners can also use the airbrakes to ensure a fast descent while still maintaining a nearly level attitude. This is mainly done for passenger comfort. The second issue that the lack of spoiler surfaces causes is an extended landing ground run. All modern jet airliners automatically deploy their spoilers on touchdown. This happens for two reasons: creating drag to slow down faster and decreasing lift to put more weight on the tires for better braking performance.

Solutions for these issues are not easily obtained without adding weight, which would eventually void the purpose of this research. Lift dumping without specific surfaces is nearly impossible. The only possibility remaining with the installed surfaces is deflecting ailerons upward. This effect would be nearly negligible but it is worth mentioning. Care should be taken with outboard ailerons however. It has been shown previously in this report that an upward deflection of these surfaces at landing speed can point the nose of the aircraft up. This would obviously nullify the desired result of firm grip on the tires.

Currently there is only one common alternative for spoiler airbraking: thrust reversers. These additions to jet engines exist in two forms, each with their own advantages and drawbacks. The first type is the target (bucket). With this mechanism two heat-resistant surfaces are put directly in the jet stream behind the engine to redirect the flow back over and under the engine. This method is be- coming increasingly rare with time because of limitations with large bypass engines. The alternative is a cascade-type thrust reverser. When these are switched on, a ring on the engine nacelle opens up to blow the bypass flow back forward around the engine. The Boeing 747 models are equipped with this type of thrust reversers, as can be seen in Figure 3.12. When thrust reversers are applied to civil aviation aircraft, it is always in combination with spoilers because of regulations. The landing ground run of an aircraft is assessed without the use of thrust reversers, for the simple reason that the aircraft needs to be able to stop adequately without engine power. Therefore this is not a truly viable alternative for spoilers. On the other hand however, one could argue that in general the landing run is shorter than the take-off run for most aircraft. It would not be wise to land on an airfield unable to take off again, so the take-off run is more critical. Nonetheless, this should be assessed for every aircraft and all circumstances (temperature, air density,...) individually.

28 Figure 3.12: Cascade thrust reverser on a Boeing 747-400

Besides the issue with the landing ground run, there is also braking in mid-air. Extending the thought process from the runway braking, thrust reversers could be an option. This is however a very dangerous approach. Firstly because the aerodynamics involved are very complex and secondly because of malfunction consequences. In case a thrust reverser gets stuck in its deployed state, this engine has to be turned off immediately and is inoperable for the remaining flight time. On a four- engine aircraft this is not too dramatic but twin engine models will have to perform an emergency landing soon after. Engines equipped with target-type reversers (which are fortunately less common nowadays) also have a considerable drag increase, even with the engine cut off. There are only a limited number of aircraft that are currently able to reverse thrust in flight, of which the Boeing C-17 "Globemaster III" is one. This aircraft was designed with this functionality for the exact same reason mentioned before: steeper descent over hazardous areas. For in-flight thrust reversal in civil aviation it seems that the drawbacks outweigh the benefits. Thrust reversal is far more radical than slightly increasing drag with spoiler surfaces. With other surfaces it would also be possible to increase the drag on the aircraft, albeit it marginally. A carefully balanced combination of various surfaces could have a net effect on only the X-axis. The most logical group would likely consist of symmetrical aileron deflection and elevator deflection for the pitching moment balance. It is obvious that the relatively small deflection of the ailerons and elevator will not nearly have the same effect as spoiler surfaces, but in near-cruise conditions the spoilers are never fully deflected. A system like this would have to be operated by the flight control computer at high frequency to make sure there are no imbalanced deflections at any time during the maneuver.

4 WEIGHT ESTIMATION

After analyzing the ﬂight performance impact of the various propositions, the next step is to assess the impact on the aircraft weight. The eventual goal of this project is to obtain an improved layout for roll control and weight saving is the part where the beneﬁts should be found. In this assessment the emphasis lies on the hydraulics subsystem. When leaving out certain control surfaces the af- fected area will not be empty, but be replaced by wing structure which would have approximately the same weight as the control surface structure. Therefore the hydraulics would be the mainly altered subsystem.

4.1. HYDRAULICSYSTEMMODEL First a rudimentary model of the hydraulic system is made. For simplicity reasons and due to lack of data and knowledge the only components considered are actuators and hydraulic lines. In reality there are numerous other components like reservoirs and pumps. However, the goal of this model is to estimate the weight loss when one specific control group is omitted and these components are generally use by more control groups at the same time, their impact is assumed to be small. An overview of the actuators and hydraulic lines for each control group of a 747-100 is given in Table 4.1. The numbers are based on the physical dimensions of the aircraft and on Figure 4.1, found in a NASA report on the reliability of the Boeing 747 control system. The distances indicated in the table are not exact but rather an approximation of the control group’s distance to the aircraft’s center point. In the breakdown the terms single feed (tandem) actuator and dual feed (tandem) actuator are used, which are linear actuators with respectively one or two hydraulic systems connected to them. The term tandem represents the fact that both sides of the hydraulic piston are supplied with hydraulic fluid. This means that both directions are actively controlled. A non-tandem linear actuator would only be able to push in one direction and would have to rely on other forces (springs, aerodynamics,...) to regain its original position. This type is basically non-existent within aircraft so the term tandem is often neglected. The control groups are broken down as follows. The inboard aileron group consists of 1 dual tandem actuator per control surface which makes two in total. The distance is measured from the aircraft’s X-axis to the inboard ailerons, following the sweep of the wing. The outboard ailerons share a similar setup, with key difference that the distance is much larger because they are located at the wing tips. The third control group located in the wing is the high-lift devices. These surface are split into 4 groups, each with a rotary actuator: inboard trailing-edge flaps, outboard trailing-edge flaps, inboard leading-edge Krueger flaps and outboard leading-edge Krueger flaps. Please note that the modeled distance of 5 m is based on no solid reference and is only an educated guess. The last wing group contains the spoilers. These 12 surfaces are all individually powered by a single feed linear actuator. Furthermore there are two ratio changer

31 Table 4.1: Actuator types and hydraulic line speciﬁcations for each control group in the weight model.

Control Group # single actuators # dual actuators # rotary actuators # lines Distance [m] Inboard ailerons 4 4 13 Outboard ailerons 2 4 28 High-lift devices 4 4 5 Spoilers 14 8 14 Central control 2 4 5 Horizontal tail 2 2 2 7 34 Vertical tail 2 1 5 34

actuators which are connected to only one hydraulic system as well. The remarkable thing of this group is the layout of hydraulic lines. Hydraulic system 1 (powered by engine 1) is not connected to any spoiler and some adjacent spoilers are connected to the same hydraulic system, together with the ratio changers. This simpliﬁes the amount of hydraulic lines to a total of 6 (three per wing).

The next group contains the central control actuators (CCA’s). These dual tandem actuators serve as control mechanism for the coupling of the ailerons and spoilers. In the early Boeing 747 models the roll control system is mainly mechanical and these actuators determine how much the spoilers deﬂect in relation to the ailerons for a given pilot stick input.

The tail holds two control groups: vertical and horizontal surfaces. Since they are all located at more or less the same distance from the planes center, the used length is identical. The horizontal tail has two separate elevator surfaces on each side, which makes four in total. The peculiar thing about these surfaces is that they do not have the same type of actuators. The outboard elevators are powered by a single feed actuator whereas the inboard elevators are connected to dual feed actuators. This is likely to be a very conscious design decision since the distance from the tip of the horizontal stabilizers to the aircraft’s hydraulic pumps is rather large. There are four separate elevator surfaces anyway so no extra redundancy was deemed necessary. On the other hand, this theory does not hold for the rudder on the vertical tail surface. This is also a split control surface but both top and bottom rudder are powered by a dual tandem actuator. Additionally, both the horizontal and vertical tail group contain a rotary actuator for trimming purposes. Since the horizontal tail is a fully adjustable stabilizer, it would be logical if its motor is substantially more powerful and heavier than the trim motor of the vertical tail, but for simplicity this difference is neglected. The exact mass of a B747-400 stabilizer trim motor was found to be 4.41 kg, coming from a manufacturer’s document [11].

The next step towards the model is assigning masses to all components in the model. According to Roskam [12] the total weight of the hydraulics in a Boeing 747-200 is 2028 kg. Furthermore, from the aforementioned manufacturer’s document [11] it is found that a rotary actuator weighs about 4.5 kg. Lastly, the linear actuators are assumed to have a mass of 10 kg and 15 kg for the single and dual respectively. With this information the last unknown is the mass of the hydraulic lines per unit distance (mline ). Reverting the equations below and solving for mline results in 2.025 kg/m. This number seems reasonable, keeping in mind that it also incorporates all unmodeled components like reservoirs etc. An overview of the component masses can be found in Table 4.2.

32 33

Figure 4.1: Flight control hydraulic system on a Boeing 747-100. [1] Table 4.2: Model component mass overview

Component Symbol Value Unit Rotary actuator mr ot 4.5 kg Single actuator msingle 10 kg Dual actuator mdual 15 kg Hydraulic line mline 2.025 kg/m

Using the hydraulics model, it is possible to calculate the approximate weight saving by leaving out certain control groups. This is according to the equations below.

m m m (4.1) hydr aulics = actuator s + lines 7 X mactuator s nr ot,i mr ot nsingle,i msingle ndual,i mdual (4.2) = i 1 · + · + · = 7 X mlines di mline (4.3) = i 1 · = The ﬁrst case, eliminating all ailerons goes as follows: Starting from a base weight of 2028 kg; subtracting the weight of two dual actuators for the inboard surfaces and two dual actuators for the outboard pair (80 kg in total); subtracting the weight of the lines (4 lines over 13 m plus 4 lines over 28 m at a weight of 2 kg/m yields 332 kg); resulting in 1616 kg. This, together with the other results are presented in Table 4.3. The full calculations can be found in Appendix A.

Table 4.3: Hydraulics system weight after modifying roll control hydraulics

Control layout Hydraulic systems mass [kg] ∆ % Regular 2028 No ailerons 1616 -412 -20.32 No ailerons with additional spoilers 1733 -295 -14.54 No inboard ailerons 1883 -145 -7.15 No spoilers 1648 -380 -18.74

4.1.1. VERIFICATION The most ideal veriﬁcation method for the model would be matching with reality. This however is not possible within the framework of this project. Torenbeek [13] provides a decent alternative. His statistical method for predicting systems weights uses the equation below to estimate the surface controls weight. Note that this systems group is not equal to hydraulics as it may contain mechanical links and electrical wiring as well. These were omitted in this project for simplicity.

W k W 2/3 [lbs] (4.4) sc = sc · TO In this equation ksc is 0.64 for jet transports with powered controls. An additional 20% needs to be added for leading-edge ﬂaps and 15% more for spoilers. Using this formula on the Boeing 747-200 with a MTOM of 377840 kg (833000 lbs) results in 3547 kg of surface controls weight. As mentioned before, the actual hydraulics weight in this aircraft is 2028 kg. This means that there is around 1500 kg of other equipment counted in the surface controls group, assuming the equation is exactly right. The important part of the Torenbeek method for veriﬁcation purposes is in the 15% addition for spoiler surfaces. Applying it to the original Torenbeek estimate would suggest a weight decrease of 463 kg. It is reasonable to assume that this 15% rule would also apply to the hydraulic part only. The assessment of no spoilers in the model resulted in a weight loss of 380 kg, which is 18.7%. This is within a reasonable margin to conclude that the model is a decent indicator of reality.

34 4.2. AIRCRAFT DESIGN INITIATOR The last step in this project is assessing the snowball effect the previously found weight savings may cause on the entire aircraft. When one component weighs less, the aircraft needs less fuel to reach its mission goal, making the aircraft even lighter, and so on. This process will be analyzed with the Air- craft Design Initiator. This software tool was first developed by R.J.M. Elmendorp [14] and has been continually modified and upgraded by several TU Delft Master students. The version that is used for this project is 2.6. It is a complex MDO (Multi-disciplinary Design Optimization) tool designed for automating the conceptual design phase of civil aircraft. The program’s flow chart is displayed in Fig- ure 4.2. The optimization loop starts with basic top-level requirements such as payload and range. From a statistical database it estimates a possible design layout which is then analyzed with performance and aerodynamics tools. If all payload, fuel and rage requirements are met the new weights are estimated. The optimization loop ends when the MTOM converges.

Figure 4.2: Flow chart of the Aircraft Design Initiator [2]

Like in every optimization program the accuracy margins are important to the end result. In the setup used for this project the weight tolerance was set to 0.5% and the range tolerance to 1%. These settings provide the best results in terms of stability, run time and accuracy all at once for the size and type of aircraft like the 747-200. To obtain the best result for this project speciﬁcally, the systems weight calculation has been altered. The regular program uses statistical relations from Raymer [15] for the weight of all subsystems, including hydraulics. See below for the equation. In this relationship N f stands for the number of functions typically performed by the controls (usually 4 to 7), L f is the total fuselage length in feet and Bw is the wingspan in feet. The result of this equation on the Boeing 747-200 is 546 lbs or 248 kg. Earlier in section 4.1 it was found that the actual weight should be 2028 kg. Firstly, this shows that Raymer estimates are not accurate for modern, large airliners which will justify some discrepancies found later in this chapter. Secondly, it demands adjustment for the purpose of this project. In further Initiator runs, the hydraulics weight will always be hard-coded as a ﬁxed number instead of a variable.

W 0.2673 N (L B )0.937 [lbs] (4.5) hydr aulics = · f · f + w With the data from the model place in the Initiator’s code, the ﬁrst task is to reproduce an aircraft that is similar to the Boeing 747-200. With the current version of the software (v2.6) it is not yet possible to design a fuselage with more than one level. Nonetheless, by setting the top-level speciﬁcations

35 of the actual aircraft (see Table 4.4) to match the speciﬁcations of a Boeing 747-200 a similar aircraft is obtained. The resulting MTOM (Maximum Take-Off Mass) and OEM (Operating Empty Mass) are displayed in Table 4.5.

Table 4.4: Top-level speciﬁcations of the Boeing 747-200

Speciﬁcation Value Unit Range 12038 (6500) km (nmi) Take-off distance 3600 m Landing distance 2130 m Cruise altitude 10668 (35000) m (ft) Cruise speed 0.86 Mach Passengers 369 - Payload 67360 kg

Table 4.5: Initiator results for an aircraft with Boeing 747-200 speciﬁcations

Parameter [Unit] Actual B747-200 Initiator result ∆ % MTOM [kg] 377842 365092 -12750 -3.37 OEM [kg] 174000 184068 -10068 -5.79

The results of the Initiator’s estimate of a 747-like aircraft are close to the actual values. If anything, the small differences (3% - 6%) can be attributed to the Raymer estimates which are not exactly on point. With the base aircraft now established, the next step is to evaluate the impact of less hydraulics weight. This assessment is done on a standard harmonic mission proﬁle. That means that the aircraft is fully loaded with payload and passengers and the fuel tanks are ﬁlled up to the MTOM is reached. This procedure does not yield the maximum range, but it is the most representative for a design case such as this.

4.3. RESULTS AND DISCUSSION

In initial results it was found that the small changes in hydraulics weight had irregular results. Some- times when decreasing the subsystem weight by 50 kg, the end result would be heavier. This incon- sistency is caused by the optimization tolerances. Instead of lowering the tolerances — which would increase the run time dramatically — a polynomial ﬁt of a set of points was chosen as solution to this problem. The graphical results can be seen in Figure 4.3 to Figure 4.5. Extracting the points coincid- ing with the values from the hydraulics model, gives proper numerical results. They are compiled in Table 4.6.

36 Figure 4.3: Maximum take-off mass vs. hydraulics mass for a Boeing 747-200 on a harmonic mission proﬁle

Figure 4.4: Operating empty mass vs. hydraulics mass for a Boeing 747-200 on a harmonic mission proﬁle

37 Figure 4.5: Fuel mass vs. hydraulics mass for a Boeing 747-200 on a harmonic mission proﬁle

The results presented above require some discussion. From Table 4.6 it can be gathered that the two more extreme cases ("No ailerons" and "No spoilers") have a greater impact than the other two. They both decrease the MTOM by about 0.5%. In the previous chapter it was determined that these two cases were not viable as they were. Although they have the most impact on the aircraft weight, it only leaves the other two cases for consideration. The case where additional spoilers were used comes close at a benefit of -0.38% MTOM. The fourth case, "No inboard ailerons", only cuts 0.19% off. The fuel saving numbers are almost linearly related with the MTOM changes: about 0.7% for the two extreme cases, 0.53% for additional spoilers and only 0.26% for leaving out inboard ailerons For the two modified cases, it is found that it would be more beneficial to remove all ailerons and to add some spoiler surfaces compared to taking out only the inboard ailerons. This is an interesting result, as it directs to justification for the philosophy of Airbus to not use inboard ailerons and instead to use more spoilers (see section 5.1). It also provides motivation for more research towards aircraft without ailerons as it proves to be feasible and even beneficial to fly with just spoiler surfaces. Focusing on the most promising case, where all ailerons are traded for some additional spoilers, there are more perspectives to look at. Besides the weight changes there is also a certain decrease in complexity and maintenance, which saves money. The fuel weight savings are an important eco- nomic aspect but after all, the change of 0.35% does point towards a dramatic impact. Looking at small modifications in the airline industry focused on fuel saving, winglets come to mind. In the early 2000’s these wing tip additions were suddenly found on all civil airliners because they could save 3-5% on regular flights [16]. This impact was so large that they were even retrofitted on old aircraft. Admittedly, the results from this project are not substantial enough to justify modification of existing aircraft, but applying the technology to newly designed aircraft seems promising. There are additional benefits to keep in mind that would surface in the design phase. In conventional commercial aircraft, the ailerons and high-lift devices take up the entire trailing edge of the wing. A new aircraft without ailerons can have high-lift devices over the entire wingspan. This brings two possible benefits, depending on the approach taken. The first is to just extend all high- lift devices over the extra space. This would yield a higher maximum lift coefficient, which in turn could increase the MTOM or shorten the take-off and landing runs. A higher MTOM is interesting for transport aircraft because it increases profit through the additional payload or harmonic range. The second approach to benefit from more trailing-edge room is to use less complex flaps over a larger span, aiming for the same maximum lift coefficient as before. The Boeing 747 model has a heavy and complex high-lift system, consisting of triple-slotted flaps and Krueger flaps on the leading

38 Table 4.6: Interpolated Initiator mass results for the Boeing 747-200 with savings percentages

Parameter [Unit] No ailerons ∆ % Add. spoilers ∆ % No IB ailerons ∆ % No spoilers ∆ % 39 MTOM [kg] 365671 -1904 -0.52 364294 -1377 -0.38 364985 -686 -0.19 363910 -1761 -0.48 Fuel Mass [kg] 113869 -553 -0.49 113469 -400 -0.35 113670 -199 -0.17 113358 -511 -0.45 OEM [kg] 184442 -1352 -0.73 183464 -978 -0.53 183954 -488 -0.26 183192 -1250 -0.68 Hydraulics [kg] 1616 -412 -20.32 1733 -295 -14.54 1883 -145 -7.15 1648 -380 -18.74 edge. Using double-slotted ﬂaps without Krueger ﬂaps over the entire wingspan has a fair chance of being lighter than the current system.

4.3.1. VERIFICATION Using complex tools like the Aircraft Design Initiator demands a form of verification. This is not to determine whether the software contains errors (which is assumed to be unlikely), but rather to eliminate the chance of a user error. The Initiator does not support a clear graphical user interface so it is not impossible that an error was made in setting up the input files and settings. Verifying the impact of hydraulics weigh specifically is not easy, but looking at it from a broader perspective offers possibilities. The hydraulics are a subgroup of the structural weight and this impact can be verified. Torenbeek [17] states the statistical relationship between structural weight and MTOM: On a subsonic transport with a range of 3000 nautical miles an increase in structural weight of 10% would result in a 7% increase in MTOM. This relationship does not apply to the material in this chapter directly. The first issue is that the Boeing 747-200 has a range of 6500 nautical miles and the second is that the structural weight only changes by 0.73% at most. Nonetheless, it shows that the relative impact on structural weight and MTOM. The fraction 10/7 1.43 (structural weight percent- = age over MTOM percentage) is very close to the fractions presented in Table 4.6: 0.73/0.52 1.40 ; = 0.53/0.38 1.39 ; 0.26/0.19 1.37 ; 0.68/0.48 1.42. Consequentially, it is likely that the Initiator was = = = used correctly.

40 5 OTHERAIRCRAFT

The results in the previous chapters were all specifically for the Boeing 747-200. The goal of this chapter is to extrapolate this knowledge to other aircraft as well. First an analysis of the types of roll control on various common airliners is presented. One specific aircraft is chosen, the Boeing 767- 200, because of the similarities in roll control. The findings from the previous chapters are applied to this aircraft and a new weight impact analysis is made.

5.1. AIRBUSVERSUS BOEING As mentioned before, the Boeing 747 models have a rather complex and redundant roll control layout. A combination of three different kinds of surfaces — inboard ailerons, outboard ailerons and spoilers — provides adequate roll performance in various situations. This is not so for all recent jet airliners, as can be seen in Table 5.1. In fact, Airbus has never produced an aircraft with inboard ailerons. Their design philosophy has stayed with the same recipe of outboard ailerons and spoilers throughout the years. Even with the A380, at this point still the largest airliner in the world, they did not find inboard ailerons to be necessary. However, it should be mentioned that the outboard ailerons stretch over approximately 25% of the wingspan and the spoilers cover more than half of the wingspan. Boeing on the other hand have stayed with inboard ailerons ever since the introduction with the Boeing 747, with the exception of the Boeing 757. A likely reason for this is the fact that it is the second smallest aircraft in their line-up. The smallest (B737) also lacks any form of inboard ailerons. In later years, with the introduction of the B777 in 1994, the design of the inboard ailerons has changed to flaperons. These surfaces can also be used as high-lift devices but still act as inboard ailerons in roll maneuvers. It was addressed before in this report that in low-speed flight the inboard ailerons have little effect on the roll rate. In wing design an important issue is the division of the trailing edge among different surfaces e.g. ailerons, flaps and engine exhaust. An inboard aileron decreases the amount of space available for flaps and thus increases the stall speed with all its consequences. It is therefore a logical choice to design a surface with dual functionality. Most aircraft in Table 5.1 are early versions of aircraft types that have gone through several up- dates. Newer versions utilize modern technologies and are often larger and heavier than their prede- cessors. Which control surfaces are used for certain maneuvers does not change in general, but the control mechanisms can be changed. For example, the Boeing 747-800 had the hydraulic spoilers and outboard ailerons replaced by fly-by-wire electrical actuators [18]. This offers improvements in terms of weight and maintenance. All other control surfaces remained fully hydraulic. Compared to other contemporary airliners this is fairly outdated. The overall trend for newly developed aircraft is full fly-by-wire. Airbus started with fly-by-wire on the first A320 and all following aircraft had full FBW

41 Table 5.1: Roll control on common jet airliners

Aircraft First ﬂight MTOM [kg] OB ailerons IB ailerons IB ﬂaperons Spoilers B737-200 1967 52 400 X X B747-100 1969 333 390 X X X B757-200 1982 115 680 X X B767-200 1981 142 880 X X X B777-200 1994 247 200 X X X B787-800 2009 228 000 X X X A320-200 1987 78 000 X X A330-200 1992 242 000 X X A340-200 1991 275 000 X X A350-800 2013 248 000 X X A380-800 2005 575 000 X X

controls. Boeing however, only started with FBW implementation in the 777 [19]. The more recent Boeing 787 was also designed with full FBW. The advantages of this modern ﬂight control system include weight saving, maintenance, redundancy management and ﬂight envelope protection.

Looking at Table 5.1, it is clear that the Boeing 767-200 is the most logical candidate to extend the knowledge found on the Boeing 747-200. The roll control systems match in nearly every way, using full hydraulic systems and the same amount of spoilers and aileron sets. There are also some striking differences, which will provide interesting results in the next section. The two foremost differences are the size and the amount of engines. The 747 is more than twice as heavy as the 767 and it ﬂies with double the amount of engines (4 vs. 2).

5.2. BOEING 767-200

From the previous section it could be gathered that the Boeing 767 is a proper candidate for extending the research of the Boeing 747. It is the only jet airliner listed with the exact same roll control layout. Additionally it is of a different weight class (less than half of the 747) so it provides a good view on the effect of aircraft size.

Regarding the flight performance of the Boeing 767, it is tough to make definite claims of what modifications would be feasible and what would have to be adjusted since there is no ready-to-use flight model of this aircraft. It will therefore be assumed that the same cases can be made as with the Boeing 747: no ailerons, no ailerons with additional spoilers, no inboard ailerons and no spoilers. The next step is to assess the weight impact of these different cases on the aircraft. Instead of designing an all-new hydraulics model for this aircraft, it is possible to scale the existing 747 model to the new 767. From Torenbeek’s formula [13] the total surface controls mass is 1990 kg. Note that this is not equal to the hydraulics mass, but it does contain the latter. In case of the 747 the actual hydraulics value (2028 kg) was 57% of the Torenbeek surface controls estimate (3547 kg). Applying the same relationship to the 767 leads to 1134 kg hydraulics mass in the Boeing 767-200.

The ﬁrst step in the Initiator assessment is again to design an aircraft that is similar to the Boe- ing 767-200. This is done in similar fashion as before, by setting the input parameters and aircraft requirements to the design speciﬁcations of the actual aircraft. They are displayed in Table 5.2.

42 Table 5.2: Top-level speciﬁcations of the Boeing 767-200

Speciﬁcation Value Unit Range 5963 (3220) km (nmi) Take-off distance 1770 m Landing distance 1463 m Cruise altitude 11887 (39000) m (ft) Cruise speed 0.81 Mach Passengers 216 - Payload 35652 kg

The results of the Initiator run with the real-life requirements and a hydraulics weight of 1134 kg results in a smaller aircraft than in reality, as displayed in Table 5.3. This was to be expected after the results of the Boeing 747-200, however the difference here is considerably larger. 21.2 % error on the OEM is not within a reasonable margin. Nonetheless, the impact of a changing ﬂight control system weight can still be assessed on this scale.

Table 5.3: Initiator results for an aircraft with Boeing 767-200 speciﬁcations

Parameter [Unit] Actual B767-200 Initiator result ∆ % MTOM [kg] 142880 125448 -17432 -12.20 OEM [kg] 80130 63141 -16989 -21.20

5.2.1. RESULTS AND DISCUSSION

In the process of evaluating the different cases for the 767-200, it was once again found that the small changes in hydraulics weight did not coincide with consistent results. Therefore a polynomial ﬁt was created to interpolate the resulting numbers. Their graphs are depicted in Figure 5.1 to Figure 5.3. The numerical values can be seen in Table 5.4.

Figure 5.1: Maximum take-off mass vs. hydraulics mass for a Boeing 767-200 on a harmonic mission proﬁle

43 Figure 5.2: Operating empty mass vs. hydraulics mass for a Boeing 767-200 on a harmonic mission proﬁle

Figure 5.3: Fuel mass vs. hydraulics mass for a Boeing 767-200 on a harmonic mission proﬁle

A comparison with between Table 4.6 and Table 5.4 yields interesting points. First thing to note is that the percentages of hydraulics are identical, which was the intention as explained earlier. Sec- ondly — and more importantly — the fuel mass percentages are also nearly identical between the two aircraft. Stuctural mass and, as a consequence, also the MTOM have a smaller percentage of decrease. These results indicate that regardless of the type of aircraft the fuel saving will typically be around 0.36% with the most promising conﬁguration, no ailerons with additional spoilers. This result is still too insigniﬁcant to be a key factor in new designs. For the other resulting numbers the same consequences and conclusions as for the Boeing 747-200 hold.

44 Table 5.4: Interpolated Initiator mass results for the Boeing 767-200 with savings percentages

Parameter [Unit] No ailerons ∆ % Add. spoilers ∆ % No IB ailerons ∆ % No spoilers ∆ % 45 MTOM [kg] 125457 -550 -0.44 125063 -394 -0.31 125263 -194 -0.15 124952 -505 -0.40 Fuel Mass [kg] 26530 -132 -0.50 26567 -95 -0.36 26615 -47 -0.18 26541 -121 -0.45 OEM [kg] 62725 -418 -0.66 62844 -299 -0.47 62996 -147 -0.23 62759 -384 -0.61 Hydraulics [kg] 903 -231 -20.37 969 -165 -14.55 1053 -81 -7.14 922 -212 -18.69

6 CONCLUSION

The goal of this thesis project was to find suitable alternatives to conventional flight control in civil aviation. The motivation behind this study was to potentially save fuel by making the aircraft lighter. Several conclusions can be drawn from the findings. They are summed up below.

First the general possibilities were evaluated for each control axis. The subject plane was a Boeing 747-200. This aircraft has a complex and exceptionally redundant flight control system. It was found that control around the Z-axis (yaw) was unsuitable for further investigation on alternative control methods with the Boeing 747-200. Besides the rudder, there are no control surfaces with sufficient yawing force on the aircraft. The engines are suitable in yaw control as they have spool-up lag and are difficult to control finely. Analyzing the control surfaces was done through simulation in a modified version of RECOVER Benchmark. The analysis of pitch control delivered the interesting result that the outboard ailerons could be a sufficient replacement for the elevators, albeit only just. The most critical maneuver during flight for pitch, the landing flare, can be achieved with full upward deflection of the outboard aileron set. This result was attributed mainly to the extraordinarily large sweep angle of the Boeing 747, making this method not viable as a general solution in civil aviation as no other aircraft has this feature.

Regarding roll control, multiple options were found because of the excessive redundancy and complexity on the Boeing 747’s wing. The aircraft uses different combinations of inboard ailerons, outboard ailerons and spoilers, depending on the circumstances, which was at the base for the alternative opportunities. The importance of spoilers during roll maneuvers was confirmed, which initiated an extensive study of several cases with different control surfaces virtually eliminated from the aircraft. The first major case was to fly without any ailerons. It was found that in cruise there were no issues of maintaining sufficient roll control but in the landing phase it was unsatisfactory. Meet- ing the legal requirement — set by EASA — was found to be impossible without ailerons. A sub-case, taking out only the inboard ailerons (thus flying with spoilers and outboard ailerons) was proposed as a viable alternative. Another modified case was proposed to meet the legal standard: adding two additional spoiler surfaces is enough to make the aircraft airworthy. Flying with only spoilers raised issues about their efficiency of deflecting airflow to obtain rolling moment. Even though it was confirmed that there is more energy loss when compared to aileron rolls, the case with additional spoilers met all requirements. The second major case to be analyzed was to fly without spoilers. This returned unsatisfactory results in regulation and landing scenarios. A solution was proposed to enlarge the outboard aileron

47 within structural limits. This proved to be a valid solution to the roll rate issue. Another possible solution was to use the rudder as well for roll control. Even though it helped to meet the requirements, it was deemed unsatisfactory because it relies on the secondary effect of the rudder through the two stability derivatives Clβ and Clr . As a consequence, it is not safe and comfortable to use in critical maneuvers close to the ground.

The second part of the research project focused on estimating the weight impact of leaving out certain control surfaces. The assumption was made that the major contributing factor to the weight loss would be the hydraulics subsystem. A hydraulics model was devised containing the main parts and lines. The extreme (unfeasible) cases of no ailerons or no spoilers resulted in a reduction of approximately 400 kg (20%) of hydraulics mass. The cases with additional spoilers and without inboard ailerons yielded a reduction of 295 kg (15%) and 145 (7%) respectively. The changes in hydraulics have a cascading effect on the entire aircraft, as was analyzed with the Aircraft Design Initiator. The results from the hydraulics model were implemented to assess the effects on MTOM, OEM and fuel mass when designing for a harmonic mission proﬁle. The two extreme cases reduce the MTOM of the Boeing 747-200 by approximately 1900 kg (0.50%) and save 550 kg (0.50%) of fuel. The modiﬁed case where two spoiler surfaces were added yielded a reduction of 1377 kg (0.38%) in MTOM and 400 kg (0.35%) in fuel. The more moderate case of only eliminating the inboard ailerons resulted in a reduction of 686 kg (0.19%) in MTOM and 199 kg (0.17%) of fuel. Lastly, the results of the Boeing 747-200 were extrapolated to the much smaller Boeing 767-200. This aircraft was chosen because of the similarities in roll control: spoilers, inboard ailerons and outboard ailerons. It was found that the effects on fuel savings with this aircraft were nearly identical percentage-wise. The structural weight and MTOM had a smaller relative reduction.

In conclusion, it can be said that the results are not impressive enough to immediately initiate renovations to the civil airline industry. It is deﬁnitely possible to control an aircraft with unconventional methods, but designing an aircraft with the proposed features would not yield enough proﬁt to justify the technical and economical risks involved.

48 7 RECOMMENDATIONS

DIFFERENT CONTROL METHODS There are still control surfaces that were not used to improve the roll rate of the aircraft because of limitations in the RECOVER Benchmark tool. Differential elevator deployment can yield promising results, as well as differential flap deployment. The latter may require changes to the actuators which would also have to be taken into account. With these two approaches included, it may be possible to make the case of flying with only spoilers airworthy without adding extra surfaces. It may even be possible to fly without any form of ailerons or spoilers, although the airbraking issue will still remain present.

MORERECENTAIRCRAFT The current state of the study is on a relatively old aircraft, the Boeing 747-200. The results may be very different on more recent aircraft. The impact on ﬂy-by-wire systems is likely to be very different than was found in this report.

MORE DETAILED MODELING The model to assess the weight improvements of the eliminated parts, was limited to only hydraulics. This can be expanded to include other subsystems as well e.g. electrical and structural. The implementation of a more detailed subsystems prediction model in the Aircraft Design Ini- tiator could yield different results, as well as a more detailed representation of the Boeing 747 with two ﬂoor levels. The overall optimization process of the software tool is also under constant development so newer versions may offer more accurate results.

Appendices

A HYDRAULICS.M

%% Input dIB = 13;%m dOB = 28;%m dCC = 5;%m dTAIL = 34;%m dSPL = 14;%m dHL = 5;%m dENG = 15;%m linesIB = 4; linesOB = 4; linesCC = 4; linesTAIL = 13; linesSPL = 6; linesHL = 4; linesENG = 4; actDTib = 2;%2 dual tandemIB actuators actDTob = 2;%2 dual tandemOB actuators actDTcc = 2;%2 dual tandem central control actuators actDTtail = 4;%4 dual tandem surface control actuators actSTtail = 2;%2 single tandem surface control actuators actROTtail = 3;%3 trim actuators actSTspl = 14;% 12 spoilers+2 ratio change actuators actROThl = 4;%2 slat rotary actuators+2 flap rotary actuators actPUMPeng = 8;%4 engine driven pumps+4 air driven pumps − − mROT = 4.5;%kg based on actual part data mDT = 20;%dual tandem linear actuator mST = 15;%single tandem linear actuator mPUMP = 13.5;%engine driven and air driven pumps − − %% Calculations linedist = dIB * linesIB + ... dOB * linesOB + ... dCC * linesCC + ... dTAIL * linesTAIL + ... dSPL * linesSPL + ... dHL * linesHL;

53 actuatorweight = actDTib * mDT + ... actDTob * mDT + ... actDTcc * dCC + ... actDTtail * mDT + ... actSTtail * mST + ... actROTtail * mROT + ... actSTspl * mST + ... actROThl * mROT + ... actPUMPeng * mPUMP; lineweight = 2028 actuatorweight;%assumption that the total is 2028 kg linepermeterweight− = lineweight/linedist; aileronsavings = dIB * linesIB * linepermeterweight + ... dOB * linesOB * linepermeterweight + ... actDTib * mDT + ... actDTob * mDT; ibaileronsavings = dIB * linesIB * linepermeterweight + ... actDTib * mDT; spoilersavings = dSPL * linesSPL * linepermeterweight + ... actSTspl * mST;

%% Results noailerons = 2028 aileronsavings noinboardailerons− = 2028 ibaileronsavings nospoilers = 2028 spoilersavings− noaileronsplus1spoiler− = 2028 aileronsavings + ... − 2*dSPL*linepermeterweight + 2*mST noaileronsplus2spoilers = 2028 aileronsavings + ... − 2*dSPL*linepermeterweight + 4*mST

54 BIBLIOGRAPHY

[1] 747 Primary ﬂight control systems: reliability and maintenance, Tech. Rep. (National Aeronautics and Space Administration, 1979).

[2] R. Elmendorp, R. Vos, and G. L. Rocca, A conceptual design and analysis method for conventional and unconventional airplanes, 29th Congress of the International Council of the Aeronautical Sciences (2014).

[3] J. F.W. Burcham, T. A. Maine, J. J. Burken, and J. Bull, Using Engine Thrust for Emergency Flight Control: MD-11 and B-747 Results, Tech. Rep. (National Aeronautics and Space Administration, 1998).

[4] C. Edwards, T. Lombaerts, and H. Smail, Fault Tolerant Flight Control: A Benchmark Challenge (Springer, 2010).

[5] L. Chen and R. Patton, Design and evaluation of FTFC scheme for sensor loss detection based on RECOVER benchmark, 2013 Conference on Control and Fault-Tolerant Systems , 649 (2013).

[6] Aircraft Accident Report: El Al Flight 1862, Boeing 747-258F 4X-AXG, Tech. Rep. (Nederlands Avi- ation Safety Board, 1992).

[7] H. Smaili, J. Breeman, T. Lombaerts, and O. Stroosma, A benchmark for fault tolerant ﬂight control evaluation, Progress in Flight Dynamics, GNC, and Avionics , 333 (2013).

[8] T. J. J. Lombaerts, Fault Tolerant Flight Control: A Physical Model Approach, Ph.D. thesis, Delft University of Technology (2010).

[9] Certiﬁcation Speciﬁcations and Acceptable Means of Compliance for Large Aeroplanes CS-25, Tech. Rep. (European Aviation Safety Agency, 2015).

[10] M. H. Sadraey, Aircraft Design: A Systems Engineering Approach (Wiley, 2013).

[11] Vickers, Fluid power, control and actuation products on the boeing 747-400, .

[12] J. Roskam, Airplane Design Part V: Component Weight Estimation (DARcorporation, 1985).

[13] E. Torenbeek, Advanced Aircraft Design: Conceptual Design, Analysis and Optimization of Sub- sonic Civil Airplanes (Wiley, 2013).

[14] R. J. M. Elmendorp, Synthesis of novel aircraft concepts for future air travel, Master’s thesis, Delft University of Technology (2014).

[15] D. Raymer, Aircraft Design: A Conceptual Approach, 2nd ed. (AIAA, 1992).

[16] B. S. de Mattos, A. P.Macedo, and D. H. da Silva Filho, Considerations about winglet design, 21st Applied Aerodynamics Conference (2003).

[17] E. Torenbeek, Synthesis of Subsonic Airplane Design (Kluwer, 1982).

[18] C. Collisson, Re-inventing the legend: the development of the 747-8, 26th Internationl Congress of the Aeronautical Sciences (2008).

55 [19] Y. C. Yeh, Design considerations in Boeing 777 ﬂy-by-wire computers, High-Assurance Systems Engineering Symposium , 64 (1998).