FIBER OPTICS FOR FLIGHT CONTROL SYSTEMS

Thesis

Submitted to

The School of Engineering of the

UNIVERSITY OF DAYTON

In Partial Fulfillment of the Requirements for

The Degree

Master of Science in Electrical Engineering

By

Bryan William Harris

UNIVERSITY OF DAYTON

Dayton, Ohio

December, 2014 FIBER OPTICS FOR FLIGHT CONTROL SYSTEMS

Name: Harris, Bryan William

APPROVED BY:

John G. Weber, Ph.D. Raul´ Ordo´nez,˜ Ph.D. Advisor Committee Chairman Committee Member School of Engineering Professor, Dept. of Electrical and Computer Engineering

Donald R. Kessler, Ph.D. Committee Member Adjunct Professor, Dept. of Electrical and Computer Engineering

John G. Weber, Ph.D. Eddy M. Rojas, Ph.D., M.A., P.E. Associate Dean Dean, School of Engineering School of Engineering

ii © Copyright by

Bryan William Harris

All rights reserved

2014 ABSTRACT

FIBER OPTICS FOR FLIGHT CONTROL SYSTEMS

Name: Harris, Bryan William University of Dayton

Advisor: Dr. John G. Weber

This paper investigates the use of fiber optic comminucation media in a digital flight control system. A number of recommendations are included allow reliable use of fiber optics in this ap- plication. A proposed fiber optic media, AFDX, is studied using dynamic, real time models. The effect of latency introduced by AFDX is studied for an F-16 aircraft model. Finally, the effect of latency and jitter are correlated using a simpler induction heater model and PID controller.

iii Thank you to my loving wife Christine, who put up with my late nights for far, far longer than it

should have taken.

iv TABLE OF CONTENTS

Page

ABSTRACT ...... iii

DEDICATION ...... iv

LIST OF FIGURES ...... viii

LIST OF TABLES ...... xi

CHAPTERS:

I. INTRODUCTION ...... 1

1.1 Origins of Flight Control Systems ...... 1 1.2 Fly by Wire ...... 2 1.3 Optical Fiber ...... 3 1.3.1 Aviation Fiber ...... 5 1.3.2 Fly by Light ...... 5

II. FIBER OPTICS ...... 7

2.1 Introduction ...... 7 2.2 Communication Wavelengths ...... 8 2.3 Mode Theory ...... 10 2.4 Modal Dispersion ...... 12 2.4.1 Graded Index Fiber ...... 13 2.5 Multiplexing ...... 14 2.5.1 Time Division Multiplexing ...... 14 2.5.2 Frequency Division Multiplexing ...... 15 2.5.3 Wavelength Division Multiplexing ...... 15 2.6 Fiber Optic Cables ...... 16 2.6.1 Physical Contact Connectors ...... 16 2.6.2 UPC/APC ...... 17 2.6.3 Non-Physical Contact Connectors ...... 17 v 2.6.4 Expanded Beam ...... 18 2.7 Losses ...... 18 2.7.1 Fiber Mismatch Losses ...... 19 2.7.2 Interface Losses ...... 19 2.8 Maintenance ...... 19 2.8.1 Splicing ...... 20 2.9 Transmitters, Receivers, Transceivers ...... 20 2.9.1 Light Emitting Diodes ...... 20 2.9.2 Lasers ...... 21 2.10 Reliability ...... 21

III. FLIGHT CONTROL SYSTEM REQUIREMENTS ...... 23

3.1 Modern Control Systems ...... 23 3.2 Unique Aircraft Requirements ...... 24 3.3 Flight Control Standards and Specifications ...... 24 3.4 AS94900 and Fiber Optics ...... 24 3.5 AS94900 FCS Requirements ...... 25 3.6 FAA AC 25.1309-1A ...... 27 3.7 Integrity ...... 27 3.8 Safe Failure Modes ...... 28 3.9 Bandwidth ...... 28 3.10 Transient Immunity ...... 29 3.11 Maintainability ...... 29 3.12 Latency ...... 30 3.12.1 Transport Latency ...... 31 3.13 Redundancy ...... 31 3.14 Determinism ...... 32 3.15 Databus Protocols ...... 33 3.15.1 MIL-STD-1553 ...... 33 3.15.2 ARINC 429 ...... 34 3.15.3 MIL-1394 ...... 34

IV. HISTORY OF OPTICAL FIBER IN AIRCRAFT ...... 35

4.1 McDonnell Douglas AV-8B ...... 35 4.2 General Dynamics F-16 ...... 37 4.3 Boeing FA-18 ...... 37 4.4 Lockheed Martin F-22 ...... 38 4.5 Lockheed Martin F-35 ...... 39 4.6 Eurofighter Typhoon ...... 40 4.7 Lockheed Martin EC 130 H Block 30 ...... 40 4.8 Boeing 777 ...... 41

vi 4.9 Boeing 787 ...... 42

V. EXPERIMENTAL FLY BY LIGHT ...... 43

5.1 Nasa F-18 ...... 43 5.2 Gulfstream 550 ...... 44 5.3 EC 135 Eurocopter ...... 44

VI. RESEARCH WORK ...... 46

6.1 Avionics Full Duplex ...... 46 6.1.1 Fiber Optic AFDX versus Legacy Avionics Busses ...... 46 6.1.2 AFDX versus IEEE 1394 ...... 47 6.1.3 AFDX Further Work ...... 48 6.1.4 AFDX Conclusion ...... 49 6.2 Latency and Jitter Effect on Control Systems ...... 49 6.2.1 SAE AS 94900 FCS Requirements ...... 50 6.2.2 F-16 Control System Bode Plots ...... 51 6.2.3 F-16 Level Flight Handling ...... 57 6.2.4 F-16 Heading Change Handling ...... 60 6.2.5 F-16 4 ft/s Altitude Change Handling ...... 64 6.2.6 F-16 30 ft/s Altitude Change Handling ...... 65 6.2.7 F-16 FCS Latency Conclusion ...... 67 6.3 Induction Heater Model and PID Controller ...... 67 6.3.1 Heater Model and Validation ...... 67 6.3.2 Physical Model Details ...... 68 6.3.3 Latency and Jitter Correlation Results ...... 72 6.4 Research Summary ...... 73

VII. CONCLUSION ...... 75

BIBLIOGRAPHY ...... 77

vii LIST OF FIGURES

Figure Page

1.1 First Fly by Wire Aircraft CF-105 ...... 3

1.2 Optical Attenuation in Glass Fibers ...... 4

1.3 Time Line of Flight Control and Fiber Optic ...... 6

2.1 EM Spectrum Used in Optical Fiber ...... 9

2.2 2D Mode Theory ...... 10

2.3 Relative Fiber Sizes for SMF, MMF and POF ...... 12

2.4 Dispersion of a Pulse in MMF ...... 13

2.5 Optical Path in Step Index and Graded Index Fiber ...... 14

3.1 Example Bode Plot ...... 25

4.1 AV-8B Harrier ...... 36

4.2 General Dynamics F-16 Fighting Falcon ...... 37

4.3 Boeing FA-18 Super Hornet ...... 38

4.4 Lockheed Martin F-22 ...... 39

4.5 Lockheed Martin F-35 ...... 39

4.6 Eurofighter Typhoon ...... 40 viii 4.7 Lockheed EC-130 ...... 41

4.8 Boeing 777 ...... 42

4.9 Boeing 787 ...... 42

5.1 Gulfstream 550 ...... 44

5.2 Eurocopter EC-135 ...... 45

6.1 F-16 Flight Control System Model ...... 51

6.2 Representative Plot of Baseline Command and Feedback ...... 53

6.3 Parametric Plot of Baseline Command and Feedback ...... 53

6.4 Parametric Plot of Time Shifted Command and Feedback ...... 54

6.5 Parametric Plot of Time Shifted and Scaled Command and Feedback ...... 54

6.6 Final Time Shifted and Scaled Command and Feedback ...... 55

6.7 F-16 Pitch Axis Open Loop Diagram ...... 55

6.8 F-16 Yaw Axis Open Loop Diagram ...... 56

6.9 F-16 Roll Axis Open Loop Diagram ...... 56

6.10 F-16 Level Flight Latency Limit ...... 58

6.11 F-16 Level Flight Unstable ...... 59

6.12 F-16 Baseline Accel ...... 61

6.13 F-16 Heading Change Latency Limit ...... 62

6.14 F-16 Heading Change Unstable ...... 63

6.15 F-16 4 ft/s Altitude Change Latency Limit ...... 64

6.16 F-16 4 ft/s Altitude Change Unstable ...... 65

ix 6.17 F-16 30 ft/s Altitude Change Latency Limit ...... 66

6.18 F-16 30 ft/s Altitude Change Unstable ...... 66

6.19 Induction Heater Under Closed Loop Control ...... 68

6.20 Induction Heater Physical Model ...... 69

6.21 Induction Heater System Physical Model ...... 69

6.22 Physical Model and Jitter and Latency Introduction ...... 70

6.23 LabVIEW Error Measurement GUI ...... 71

6.24 Best Fit of Latency and Jitter On PID Control System ...... 73

x LIST OF TABLES

Table Page

2.1 Common Wavelengths for Optical Communication ...... 8

6.1 Comparison of AFDX Timing Parameters vs. Other Busses ...... 48

xi CHAPTER I

INTRODUCTION

1.1 Origins of Flight Control Systems

Heavier-than-air aircraft have been flying for over 100 years since the development of the Wright

Flyer. Work on aircraft began long before this. Sir George Caley is reported to have flown a winged coach as early as 1852[1]. Experimentation with hang gliders had been conducted successfully by the Wright Brothers and others. The Wright brothers had many competitors with larger or more powerful aircraft[1, 2]. The 1903 Wright Flyer was special because it was controllable[2]. Stable,

3-axis flight control is critical to powered flight. Flight control systems (FCS) have progressed tremendously over the past century and have greatly increased in both performance and complexity, but stability remains a chief concern[1, 3, 4, 5, 6].

The Wright Flyer used cables and levers to move flight control surfaces. As planes became larger and faster, aerodynamic loads increased. When mechanical levers were no longer sufficient to overcome aerodynamic loads, hydraulic actuators were introduced to help move flight surfaces.

Aircraft continued to progress in both performance and complexity and Flight Control Computers

(FCCs) were introduced to help maintain aircraft control. Fly by Wire (FBW) decoupled the pilot from from the individual control surfaces and allowed the aircraft to reach higher levels of control performance than a pilot alone could mechanically provide. The low level responsibility for flight 1 control was transferred first to analog and eventually digital computers. Eventually, many special purpose computers were put into airplanes for every conceivable function. Digital signal buses were developed to communicate amongst the many computers to pass digital data, and fat, round coaxial cables were installed for the radars. As aircraft became more complicated, they began to fill up with copper wires[7].

Modern aircraft contain many miles of copper wire to link the computers, sensors, avionics, displays, and even entertainment systems. The size and weight of copper communications cables now make up a significant portion of the aircraft[8, 9]. Fiber optics provide an opportunity to take back some of that size and weight and to enable much higher bandwidth required for the next generation of FCS[10, 11]. Although fiber optics have already been introduced into aircraft and have been flying in aircraft for over 30 years, no production aircraft has yet to utilize fiber optics for its primary FCS despite the advantages of fiber media over copper[12]. Because fiber Optics provide distinct advantages over copper wire they should be considered for use in the modern, digital, Fly by Wire flight control system.

1.2 Fly by Wire

The Avro Canada CF-105, shown in Figure 1.1[13] was the first aircraft designed and flown with a FBW flight control system. It used an analog computer for flight control as early as 1958[14, 15].

The first production aircraft to use FBW was the Concorde supersonic passenger airliner, introduced in 1976[15, 16]. The early F-16 used analog computers to help maintain stability in flight and improve maneuverability[15, 17, 18]. The first digital FCC was developed by NASA for the space shuttle in the late 70s[15]. The F-16’s analog FCC was eventually upgraded to a digital system in the C/D variant, block 40 in 1988 and 1989 respectively[19]. The first all digital commercial FBW airliner was the Airbus A320 introduced in 1984[15]. Boeing first introduced a FBW aircraft in

2 Figure 1.1: First Fly by Wire Aircraft CF-105

1995 with the 777[15, 20]. All modern fighter aircraft, bombers such as the B2, cargo planes such as the C-17, many Unmanned Aerial Vehicles (UAVs) and many modern airliners have FBW FCS and digital FCCs[18, 21, 22, 23, 24, 25, 26].

1.3 Optical Fiber

The first practical glass optical fibers were produced in 1970 by Corning with 20dB/km atten- uation. By 1977, telephone field trials were performed with 50µm and 62.5µm fibers[27]. Single mode fibers were developed for transatlantic cables and telecommunications companies built na- tional fiber backbones in the early 1980s[27]. Modern, single mode fibers (SMF) can transmit 100s of gigabits of data per second over a single 5-8µm fiber with 0.2 to 0.3 dB/km losses[28]. Figure

1.2[28] shows the theoretical attenuation of light in glass versus wavelength.

More recently, plastic optical fibers (POF) have emerged which have some but not all the ben- efits of glass fibers. The signal attenuation in POF tends to be about an order of magnitude higher

3 Figure 1.2: Optical Attenuation in Glass Fibers

than glass fibers which limits the useful bandwidth and distance of links using these fibers. How- ever, POF can be much easier to work with than glass fiber and enables lower cost for short, low bandwidth links[28].

4 1.3.1 Aviation Fiber

Fiber optics have been in aircraft for many years. In some cases, fiber has been used in mission- critical systems. The A-7 ALOFT program demonstrated that fiber could be used on aircraft in

1977. The AV-8B had a mission-critical fiber link since 1981[29]. The 777 has an extensive fiber network for both in-flight entertainment and avionics since 1995[20, 24]. The Eurofighter Typhoon uses SAE AS1773 (1773) buses for avionics systems which transmit mission-critical information over fiber optics[26, 30]. Finally, the F-22 and F-35 have extensive, mission-critical fiber optic networks[31, 32, 33]. Despite numerous examples of mission-critical fiber optic links, no known aircraft has yet used a flight-critical, fiber optic link.

1.3.2 Fly by Light

The technologies of FBW and fiber optics have advanced in parallel as summarized in Figure

1.3. Flight control developments are shown in blue and developments involving fiber optics are red.

FBW has become the standard for aircraft flight control and fiber optics are the media of choice for high performance communications. A number of attempts have been made to combine the two and create Fly by Light (FBL).

The NASA/Navy Fiber Optic Control System Integration (FOCSI) program explored the inte- gration of fiber optics into a flight control system, including digital communications and passive optical sensors in the early 90’s[34, 35]. Three FBL Tech Transition Symposia were held in the mid

90’s in an attempt to popularize the idea of FBL[36, 37, 38]. HR Textron built an optically con- trolled aileron actuator which was flight tested at NASA LARC in 1997[39]. Gulfstream performed similar experiments with a single secondary control surface on a G550 in 2008[40]. Gulfstream used a novel technology which places electrical to optical converters inside a standard size electri- cal back-shell. The French Rafale and Mirage reportedly had experimental versions in which fiber

5 was used to transmit pilot commands to the FCC in the mid 80’s[41, 42]. Finally, there is a version of the Eurocopter EC135 which replaces its entire data bus with a fiber optic backplane[43, 44].

The Active Control Technology demonstrator/Flying Helicopter Simulator (ACT/FHS) is used to simulate other helicopters by actively simulating the handling qualities of those aircraft in real time.

This helicopter system is perhaps the most mature example of a Fly by Light flight control system and has been in use since 1992.

CF-105 FBW ConcordeF-16A/B FBWShuttle FBWA320 DFCC/FBW DFCC/FBWF-16C/DC-17 FBW/DFCC777 FBW/DFCC DFCC/FBW

Glass-Clad Fiber Corning SMF AV-8B Link 777 BusNASA F-18Eurocopter FBLEurofighterF-22 FBL Bus Bus 1950 1965 1980 1995 2010

Figure 1.3: Time Line of Flight Control and Fiber Optic

6 CHAPTER II

FIBER OPTICS

2.1 Introduction

Fiber optics are long cylindrical pieces of glass or plastic which transmit light through the prin- ciple of total internal reflection. Light is injected at one end by a light emitting diode (LED) or a laser and received by a photo detector at the other end. Fibers allow the transmission of large amounts of information over very long distances as compared to electric wires. Fibers are immune to many sources of electro-magnetic noise which affect electric wires, are much lighter, are not susceptible to electrical arcs do not cause ground loops. In telecommunications, fiber optics have proven to be extremely reliable when compared to copper wire[29, 45].

There are typically many layers to a modern fiber but the key parts include the core, cladding and jacket. The core, typically measured in µm, is the tiny glass center of the fiber which carries light.

The cladding is made of a slightly different glass with a lower index of refraction, which surrounds the core and keeps the light inside the core. Light which enters the cladding is rapidly attenuated.

The jacket protects the core and cladding and is not glass but typically a polymer material.

7 2.2 Communication Wavelengths

A schematic of communications wavelengths and their relation to visible light is shown in Figure

2.1. The light used to transmit data in fibers is typically a longer wavelength than visible light, though shorter wavelength transmitters will sometimes appear red. Humans can see light from about 390 nm (purple) to 750 nm (red). Communications wavelengths are typically in the infrared which is about 700 nm to 3000 nm. Table 2.1[46] on page 8 is a summary of common wavelengths used to transmit data on optical fiber.

Table 2.1: Common Wavelengths for Optical Communication

Band Wavelength [nm] NIR 800-900 O band 1260-1360 E band 1360-1460 S band 1460-1530 C band 1530-1565 L band 1565-1625 U band 1625-1675

8 Figure 2.1: EM Spectrum Used in Optical Fiber

9 2.3 Mode Theory

A simplified picture of light in a fiber can be seen in the two dimensional, parallel mirror waveg- uide theory. Geometric paths which interfere constructively with themselves tend to propagate. If light were pictured as a wave, then this occurs when the peaks and troughs of successive waves line up. This happens when the parallel paths traveled by a light wave differ in length by an integer multiple of the wavelength as in Equation 2.1 and Figure 2.2.

Figure 2.2: 2D Mode Theory

λ d sine(θ ) = m ; m = 1, 2, 3 ... ; sine(θ ) ≤ 1 ⇒ m ≤ 2 (2.1) m 2π m π

With parallel mirrors, all light is assumed to be reflected no matter the angle. With dielectric waveguides such as plastic or glass, light can escape when its incidence angle is above a certain crit- ical angle θc. A common measure of fiber to retain light below the critical angle is called numerical aperture (NA). The NA is dependent on fiber geometry as described in Equation 2.2.

2 2 NA = sine(θc) = (n2 − n1) (2.2)

10 The critical angle is determined based on the difference in index of refraction inside and outside the waveguide. Substituting this into Equation 2.1 gives Equation 2.3.

d sine(θ ) ≤ sine(θ ) ⇒ sine(θ ) ≤ 2 NA (2.3) m c m π

The cylindrical geometry of fibers allows for much more complex mode shapes which are not easily visualized. Waves can propagate not only with different angles, but can coil around the inside the fiber and a number of different counter-intuitive shapes are possible. What is important to understand is that a mode is a specific shape of light wave that is supported by the fiber geometry and propagates down a fiber with little or no interaction with other modes. A rough number of propagating modes can be found from the Equation 2.4 where M is the number of propagating modes and V is a dimensionless parameter which describes a fiber’s ability to support multiple modes of light waves.

d 4 d2 For Large V; V = 2π NA; M ≈ V 2 + 2 ⇒ M = 16 NA2 + 2 (2.4) λ n2 λ2

The number of modes is proportional to the diameter of the fiber and inversely proportional to the wavelength. There is a special geometrical case where only a single mode is possible when the

fiber diameter is very small as compared to the wavelength. A fiber small enough to propagate only a single mode is called Single Mode Fiber (SMF). The requirement for fiber to be SMF is found in equation 2.5. Since all light traveling down a SMF takes the same path, there is no modal dispersion.

This lack of dispersion allows extremely short pulses to be resolved after traveling long distances and thus allows extremely high bandwidth. When V is less than 2.405, multi-mode transmission does not occur.

11 V < 2.405 → Single Mode Fiber (SMF) (2.5)

SMF have a dramatically smaller core size than MMF and POF has no separate cladding section at all. This size difference is shown in graphically in Figure 2.3.

Figure 2.3: Relative Fiber Sizes for SMF, MMF and POF

2.4 Modal Dispersion

Longer wavelengths are used for communication partly because they reduce the number of modes that propagate on a fiber and minimize dispersion. Light is an electromagnetic wave and each mode is a specific shape of light-wave which will propagate down a fiber. Each mode in turn has a slightly different propagation velocity down a fiber which leads to modal dispersion. A pulse

12 of light which is initially sharp and well defined will decrease in amplitude and spread out as it travels down a fiber as shown in Figure 2.4[47].

Figure 2.4: Dispersion of a Pulse in MMF

2.4.1 Graded Index Fiber

Graded index fiber is a way to reduce modal dispersion in multi-mode fiber (MMF) without reducing the size of the fiber core. This is done by gradually changing the index of refraction of the core and clad material. In graded index fiber, the modes curve at the interface between the core and cladding instead of bouncing sharply. This reduces the differences in the path lengths traveled by different modes and thus reduces dispersion as shown in Figure 2.5.

13 Figure 2.5: Optical Path in Step Index and Graded Index Fiber

2.5 Multiplexing

Fiber optics support enormous data transmission speeds (bandwidth), which makes it possible to transmit many optical signals down a single fiber. Transmitting multiple distinct signals down a single channel is called multiplexing. Three generic methods are Time Division Multiplexing

(TDM), Wavelength Division Multiplexing (WDM) and Frequency Division Multiplexing (FDM).

2.5.1 Time Division Multiplexing

The most straight forward method of multiplexing is TDM. Using this method, only a single sig- nal is transmitted, one at a time using a round-robin or other time-sharing algorithm. This method requires only a single transmitter and receiver and no special optical hardware to combine or sep- arate the optical signals. However, the bandwidth of such a system is limited by the electronic transmitter and receiver and the full theoretical bandwidth possible on the fiber is not realizable with current technologies. Moreover, the requirement for a single signal at a time can add latency to the signals which may not be desirable.

14 MIL-STD-1553 (1553) and 1773 bus are examples of buses which use TDM to share one bus among many senders and receivers. These buses transmit within preallocated time slices which are guaranteed to be available at a specific interval. Serial buses such as IEEE 1394b (1394) can also use

TDM to allow multiple senders and receivers. Packet switched networks, such as the IEEE 802.3

(Ethernet) are generally not considered TDM devices, but share some characteristics with TDM systems. For instance, while a full duplex ethernet device can transmit and receive simultaneously, the device cannot simultaneously send or receive two messages at once.

2.5.2 Frequency Division Multiplexing

FDM uses carrier signals with various frequencies with some sort of modulation to transmit data. This is the method used in classical Radio Frequency (RF) transmission. This method can utilize standard RF tuning devices. This method is used to transmit raw RF signal content over long distances after conversion from electricity to light. Unlike TDM, this allows multiple signals to be transmitted simultaneously. Signals transmitted using FDM can interfere with each other because the information is carried by the light amplitude. Also, modulation and demodulation hardware tends to limit the bandwidth which can be transmitted below the theoretical bandwidth possible on

fiber[48].

2.5.3 Wavelength Division Multiplexing

WDM uses the frequency of the light itself to divide signals rather than a carrier signal. Like

FDM, this allows multiple signals to flow simultaneously down the same fiber. Unlike FDM, multi- plexing and de-multiplexing can be accomplished by passive optics which can have the same band- width as the fiber itself. Moreover, a larger number of signals can be transmitted than with either of the other two methods allowing the highest bandwidth of the three methods. Signals multiplexed with WDM do not tend to interfere with each other because the are separated by wavelength and are

15 not necessarily dependent on light intensity. Telecom equipment almost exclusively uses WDM and the number of signals transmitted down a single fiber is constantly increasing as narrower bands of wavelength can be resolved.

There has been considerable interest recently in WDM for avionics applications[49, 50, 51]. In addition to the increased bandwidth, WDM enables the strict segregation of networks with mixed criticality. This important feature is needed to guarantee that less critical systems sharing the same optical network cannot interfere with the FCS. For instance, the entertainment systems could share

fiber with the flight control system without the ability to affect flight controls.

2.6 Fiber Optic Cables

There are too many types of fiber optic interconnects to conduct an exhaustive review of all cables and connectors. However, connectors can be classified in a number of general ways. One such classifier is the number of fibers contained in a single connector. Another is whether or not the the fiber end faces physically touch inside the connector. The geometry of the physical contact surface is another classifier. Finally, U.S. military connectors all include round screw-on back-shells which are designed to handle the rough treatment of a military environment.

2.6.1 Physical Contact Connectors

Physical contact (PC) connectors are the most common type. In PC connectors, the end facets of fibers actually touch each other and are designed to never lose contact once mated. Common industrial PC connectors include the ST, FC, LC, and MT types. MIL STD 38999 connectors and nearly all telecom connectors are PC.

These connectors have the lowest interface losses due to their necessary tight tolerances and lack of an air gap. PC connectors are utilized to minimize signal loss, but are sensitive to contamination and damage. They are very reliable when left connected, but typically have limited mate-demate 16 cycle ratings. Repeated mating cycles will tend to reduce the performance of a fiber link, particularly if the end facets are not cleaned carefully prior to mating. Rated mating cycles for PC connectors are nearly always in the low hundreds. Microscopic contaminants, invisible to the eye, can render a

PC connector inoperable and can permanently damage a PC connector when mated.

2.6.2 UPC/APC

Ultra Physical Contact (UPC) and Angled Physical Contact (APC) connectors are slight mod- ifications to the contact surface and confer additional performance over standard PC connectors.

APCs have a slight angle to the contact face which greatly reduces the reflected light from the inter- face, by up to 30 dB. This reduces the noise on a link due to reflected light and maximizes available bandwidth. UPC connectors have a dome shaped contact surface which reduces losses by about 10 dB versus a standard PC connector.

2.6.3 Non-Physical Contact Connectors

Non-physical contact (NPC) connectors fiber faces do not actually touch and can offer increased reliability for connectors which must be demated often. NPC connectors can reduce or eliminate some weaknesses of PC connectors at the cost of increased link losses. Because the faces do not touch, mating cycles will not tend to cause damage to the fiber ends. Faces may still become dirty, but will not tend to be permanently damaged by contaminants. If link quality is reduced due to contamination, links can typically be returned to the original state through cleaning rather than re-termination or re-polishing.

NPC connectors have been used successfully in aircraft mission systems. The early serial data link installed in the Harrier made us of a NPC SMA 904 connector[29]. Likewise, the Arinc 636 bus in the Boeing 777 makes use of NPC connectors to ease maintenance and increase mating durability in harsh environment locations[20].

17 2.6.4 Expanded Beam

Expanded Beam (EB) connectors address the sensitivity to losses in connectors due to imperfect geometry, contamination, and air gaps. This is accomplished by magnifying and collimating the light coming from one fiber and then refocusing the light back into the next fiber. In this way, the interface between connectors becomes physically larger. EB connectors may be either PC or NPC type, but are distinguished by the larger (expanded) interface area. Small contaminants which might block an entire fiber core may not even cause a problem for an EB connector. Even scratched or dirty EB connectors may still be usable.

EB connectors are sometimes used in fibers which must transmit extremely high optical powers, such as in optically powered amplifiers. One consequence of expanding the beam is that the optical power density at the surface interface is reduced. In the case of optically powered amplifiers, the optical power can be high enough to physically damage the fiber transmitting the light. Contami- nants which absorb light can be heated until the fiber burns. Once the glass is darkened the damage can cascade back toward the optical source and destroy the entire fiber. The risk for contamination is greatest at the mating surface of the connector. EB connectors reduce this possibility by reducing the optical power density at a connector interface to a level which cannot physically damage the

fiber[52, 53].

2.7 Losses

There are a number of ways in which optical power can be lost in a fiber optic connector. Intrin- sic losses are caused by the connection of fundamentally different fibers and cannot be improved by

fiber connectors. These losses include those caused by fiber diameter mismatch and numerical aper- ture (NA) mismatch. Extrinsic losses are typically controlled by the connector and include angular and axial misalignment and the size of any air gap.

18 2.7.1 Fiber Mismatch Losses

A diameter mismatch typically causes an optical power loss in only one direction and is caused when transmitting from a larger diameter core fiber into a smaller diameter core. NA is a function of the angle of acceptance of a given fiber. This is also a unidirectional loss and will occur when transmitting from a larger NA fiber to a smaller. These losses cannot be improved by any connector type, including an expanded beam connector.

2.7.2 Interface Losses

Interface losses are a direct result of the connector interface and can be affected by changing the connection type. The maximum misalignment of a given connector is controlled by the tolerances used in the manufacture of a given connector as well as its proper assembly and workmanship. As- sembly of fiber optic connectors and cables is non-trivial and typically requires specialized training and equipment. Moreover, defects in fiber cables are often invisible to the eye and imperfect cables cannot always be visually identified. Thus interface losses must be carefully controlled or verified for proper operation of optical links.

2.8 Maintenance

Maintenance is not typically required for terrestrial fiber as long as it is left alone. Required maintenance is typically limited to the replacement of accidentally broken fibers and cleaning of

fiber end faces prior to mating. This requirement for cleaning end faces is a key difference between copper and fiber links. Repair and cleaning of optical fiber requires specialized equipment and training. The cleanliness of fiber end faces is very important, particularly as the core sizes get smaller. Microscopic contaminants can block enough light on an fiber end-face to prevent a fiber link from operating and hard particles trapped in mating surfaces can permanently damage end

19 faces. Termination of cables requires special equipment which can be difficult to use outside of a laboratory environment. Maintenance has been a stumbling block for the adoption of flight-critical

fiber optic links[54].

2.8.1 Splicing

Repair of broken fibers is particularly difficult in single-mode glass fibers. Single mode fibers have a core size approaching a few µm in diameter and therefore require sub-micron positional accuracy to achieve low loss through a splice. Much work has been done on splicing over the years, but the process of splicing glass fibers remains a technically difficult process. Plastic optical fibers are easier to splice or terminate because their large core sizes reduce the need for precise positioning.

In some cases, POF can even be terminated without special equipment.

2.9 Transmitters, Receivers, Transceivers

There are a few ways to transmit and receive light in an optical fiber. Common light transmit- ters include the Light Emitting Diode (LED) and Light Amplification by Stimulated Emission of

Radiation (laser) device. Either of these devices may be modulated directly by turning them on and off or externally modulated to encode data on the light signal. Receivers include photo-diodes, photo-transistors, and photo-multipliers. Each of these devices have advantages and disadvantages which will be discussed below.

2.9.1 Light Emitting Diodes

LEDs are the cheapest type of transmitter and are typically directly modulated. These devices transmit a relatively broad spectrum of light compared to lasers. This broad spectrum and a large cone of light produced makes it difficult to couple all of the light from an LED into a fiber. These

20 devices also tend to support lower modulation rates and have lower optical power than lasers. How- ever, LEDs are easier to power than lasers and have a wide range of operating current and voltage allowing for very simple driver circuits.

2.9.2 Lasers

Lasers are more expensive that LEDs, but are becoming common due to advances such as CD players and DVD and Blue-Ray disc players. They can be directly modulated, but are often modu- lated externally. They have a narrow range of operating voltage typically, and need a more complex driver circuit than an LED. Lasers typically have a very narrow frequency range and are often re- ferred to as monochromatic. The are also coherent, which means that all light transmitted has the same phase.

Lasers can couple more power into a single-mode fiber than LEDs because of their narrow spectrum. External modulators tend to have higher bandwidth than direct modulation. For these reasons, lasers with external modulation are most often used for long haul telecommunications.

However, it is unclear if these devices are the best fit for aircraft flight control because links tend to be short and have modest bandwidth requirements.

2.10 Reliability

One question when replacing electronic links with fiber is whether the reliability of the links might suffer, particularly in the harsh aircraft operating environment. A number of studies have been conducted on the reliability of fiber optic links. One study by Rome Air Development Center compared the reliability of copper telecom links to fiber optics. Failures were so rare in this study that obscure failure modes were included such as “Chewing by Rodents” and “Hunters Shooting at

Birds”[29].

21 An aircraft has a number of stresses which do not exist in telecom installations. The temperature extremes and thermal shocks are far greater in an aircraft than a static installation. Moreover, the mechanical vibration and shock seen in an aircraft is not present in a static outdoor environment.

Finally, an FCS demands higher reliability than a telecom link which does not typically result in a loss of life in the case of a failure.

More work needs to be done to demonstrate the reliability of fiber optics in aircraft. At present, there is little published information on the reliability of fiber optics in a harsh aircraft environment.

Some of this test data may already exist because of the increased use of fiber in avionics systems, but little has been published.

22 CHAPTER III

FLIGHT CONTROL SYSTEM REQUIREMENTS

3.1 Modern Control Systems

Flight control systems (FCS) have grown in complexity and performance over the years. Com- ponents of a FCS include the pilot interface (both display and control), computers, data links, me- chanical actuators and flight control surfaces. Many of these flight-critical components could benefit from fiber optic links. Detailed standards have been developed to ensure that FCS will perform as expected in all circumstances. Such standards impose unique performance requirements and relia- bility constraints on any data link including fiber optics.

Modern FCS employ sophisticated and powerful digital computers to maintain aircraft control.

The flight control computers and their associated links have a number of unique requirements in- cluding very high integrity in harsh environments, moderate bandwidths, immunity to transients, maintainability, error detection/correction, intrinsically safe failure modes, and low latency. In- tegrity requirements nearly always lead to the use of multiply redundant systems in manned FCS and push the limits of what is physically practical for modern electronic devices.

23 3.2 Unique Aircraft Requirements

The FCS has a number of requirements not typically found in other applications. The flight control system is a safety critical, dynamic, real-time system, which is increasingly necessary to maintain control. For some systems, the aircraft is so dynamically unstable without FCS augmen- tation, that any momentary failure of the flight control system could result in the destruction of the aircraft. FCS require a technologically difficult combination of determinism, integrity, fault tolerance, and intrinsically safe failure modes.

3.3 Flight Control Standards and Specifications

There are a number of publicly available standards documents which pertain to US flight control systems. SAE AS94900 is a former military standard which provides design guidance for FCS[5].

FAA Federal Aviation Regulations (FAR) parts 25 and 27 are the U.S. federal guidelines which pertain to flight control[55]. The MIL-F-8785 gives guidance on the handling qualities which should be provided by FCS. Finally, SAE AIR 4235A discusses actual flight control systems in aircraft and demonstrates some of the extraordinary measures designers have taken to ensure flight safety[56].

3.4 AS94900 and Fiber Optics

AS94900 is the standard for military flight control systems. This document lists specifications for hydraulics, three for hydraulic servovalves alone, actuators, fittings, power units, forgings, bear- ings, etc. However, only AS50881 and 1773 deal with fiber optics[7, 57]. AS50881 is the standard for aircraft electrical wiring and has a recently added section on fiber optics. 1773 is an upgrade from a copper avionics bus to a fiber bus. Other than the title of SAE AS1773, fiber optics are not even mentioned in AS94900. Fiber is very different than metal wire despite a similar outward appearance, but is largely treated like wire in applicable standards[5, 7]. For instance, copper wires

24 are not easily broken by minor impacts or bending, while fibers are. Furthermore, the cleanliness required by fiber optic end faces is in a completely different class than electrical wires. Dirty fiber optic end faces may cause optical links to perform poorly or more likely not function at all. A large percentage of the replaced fiber optic cables used in avionics can be attributed to dirty fiber end faces[58, 54]. Clearly, standards need to be written for fiber optics in FCS. Recent progress has been made in aerospace fiber standards, but the lack of applicable standards has been a barrier to adoption of fiber in the aerospace industry[59, 58, 12].

3.5 AS94900 FCS Requirements

There are a number of specific, quantitative requirements in AS94900 regarding the handling and performance of aircraft. These include generic phase margin (PM) and gain margin (GM) requirements along with specific maneuvers. The requirements vary with the type of aircraft, but this section will outline the requirements for the highest performance aircraft.

Figure 3.1: Example Bode Plot

25 The aircraft flight control system is required to have greater than 6 dB gain margin and 45° phase magrin. These quantities are typically illustrated on a Bode Plot as shown in Figure 3.1. Gain margin is the system gain at the frequency where the phase lag passes 180°. At this point, the system passes from a negative feedback to a positive feedback system. The gain margin requirement of 6 dB attenuation guarantees that inputs at these positive feedback frequency are highly attenuated.

Phase margin is 180° minus the phase at which the system drops below unity gain. The phase margin requirement guarantees that signal content with frequencies within 45° of causing positive feedback are always attenuated to some degree.

FCS cannot cause pilot induced oscillations. The FCS cannot flutter or diverge. This seems sensible. Each actuator must have 6dB gain and 45° phase margins. These margin requirements match the aggregate system requirement.

In steady flight there are flight smoothness requirements. The aircraft must not produce more than 0.05 g vertical or 0.02 g lateral acceleration. It must not sideslip more than 1° or have more than 0.5° of pitch angle. The FCS must control altitude to within 30 ft, heading to within 0.5° and speed within 0.01 mach or 2% of indicated speed whichever is greater.

The aircraft must remain stable during a momentary heading change. This could be caused by turbulence or an accidental control input by the pilot. During a heading transient, the aircraft must not overshoot more than 0.5°.

During a heading change, the maximum deviation is 1° with a 2.5° overshoot. The roll rate must be less than 5°/sec. During a bank, the sideslip angle must be less than 2° and the maximum lateral g force is 0.03 g. While rolling, the maximum lateral acceleration is 0.05 g. The maximum pitch deviation is 1° and the settling time must be less than 3 seconds. During a heading change, the speed must not change more than 10 knots or 2% of indicated speed, whichever is greater.

26 During an altitude change there is a maximum of 0.5 g normal acceleration, and the speed must not change more than 10 knots or 2% of indicated speed. At a ramp rate of 4 feet per second, there must be less than 20% overshoot and less than 175% pitch overshoot. During a 30 foot per second ramp, the deviation must be less than 5° per second from the commanded pitch. A 5 foot per second siulated gust must not cause more than 1 ft altitude error or 40% pitch overshoot.

3.6 FAA AC 25.1309-1A

The FAA published an Advisory Circular for flight critical systems in 1988 entitled, ”AC25.1309-

1A, System Design and Analysis” (FAA AC 25.1309)[60]. This standard contains many of the same requirements at AS94900. This FAA document specifies that a safety critical flight control system have less than 10−9 probability of failure per flight hour. It also specifies that flight control systems should utilize redundant and backup systems to ensure that no two component failures should result in an unsafe condition. It also requires inherently safe failure modes, tolerance of manufacturing errors, failure warnings, isolation of systems to prevent cascading failures, and the availability of self-check systems.

3.7 Integrity

AS94900 requires an FCS integrity of up to 10−8 failures per flight hour. In addition, the failure of any component must either be ten times less likely (up to 10−9) or must not degrade the aircraft performance. This requirement is very strict and even states that a failure must not lead to erroneous or misleading information to the pilot which could result in an incorrect pilot input. An off-the-shelf

fiber optic link, cannot meet this integrity requirement, but neither can any other type of data link.

In telecom applications, fiber has been shown to be more reliable than copper wire, but this has not yet been demonstrated for an aerospace environment[29].

27 Integrity is concerned with more that just breakage of components. An FCS must also guarantee that undetected errors do not occur. Standard error checking algorithms in communication systems are typically not sufficient to ensure 10−9 integrity. The 32-bit CRC algorithm used by 1394 and

Ethernet to detect transmission errors only provide about 10−7 integrity. In other words, there is a

10−7 chance that a bit error will go undetected leading to data corruption. Additional layers of error

checking need to be added to these protocols to ensure adequate integrity for a FCS.

3.8 Safe Failure Modes

Flight control systems are required to undergo a rigorous analysis which looks at all predicted

failure modes and guarantees that no single failure or combination of failures will cause the airplane

to crash or even become unstable. FAA AC 25.1309 even requires that FCS be analyzed assuming

that any single failure has already occurred. This guarantees that the system will be fully operational

“(Fail-Op)” if any single failure occurs. The FAA also requires that a second failure must also leave

the system fully operational “(Fail-Op, Fail-Op)”. This means that flight critical data links must be

at least triple redundant with some way of determining in real time which link is bad and which is

good. In practice, this leads to triplex or quad redundant FCS.

The same error detection and redundancy solutions for electrical links should be applicable to

fiber links. This redundancy requirement actually multiplies the weight and size benefit of fiber

optics over electrical links. A broken digital electrical link looks very much like a broken digital

link over fiber. In fact, as discussed previously, the dielectric nature of fibers inherently prevents

common electrical problems such as ground loops, crosstalk, and EMI susceptibility.

3.9 Bandwidth

The moderate bandwidth requirements of a flight control system would not typically be chal-

lenging for telecom equipment. Legacy busses such as 1553 and ARINC 429 have low bandwidths, 28 1 Mbit/sec and 100 kbit/sec respectively. Moreover, the strict TDM used in 1553 further limits its bandwidth. The 1553 bus suggests update rates of 400 Hz or less and not more than 3 kHz.

Large core, plastic fiber such as that which is used in home audio equipment or home networks does have limited bandwidth. A typical POF tends to have about the same bandwidth as a similar

Cat5e networking cable or 100 Mbit/sec for up to 100ft. This rate would be adequate for nearly all current FCS given the speed of legacy systems, but bandwidth considerations could limit the use of

POF in a future flight control system.

3.10 Transient Immunity

The immunity to electrical transients could be a challenge to fiber links in an FCS. Aircraft power systems can have extremely large transients, particularly military aircraft[61]. Mil-std-704E allows transient spikes of up to 425 V on the 270 VDC bus and 50V on a 28 VDC bus. These transients would not affect the fiber itself, but fiber transceivers could be interrupted or damaged by such transients[62]. Laser sources are particularly sensitive to over-current spikes and would need to be protected against such conditions.

3.11 Maintainability

Maintainability is a challenge for aerospace fiber optics[12, 63, 29]. Even in a terrestrial envi- ronment, the repair of fiber can be more complicated that the copper it replaces if one considers a single fiber link vs a single copper link[10]. However, when accounting for the fact that a fiber can replace many electrical connections, it may be that fiber will eventually be more maintainable than the aggregate of the copper wires it replaces.

Aircraft maintenance involves frequent disconnection of wiring harnesses. A typical PC con- nector used in a telecom is rated at 250 mating cycles which would not be adequate for a avionics

29 environment. Standard aircraft maintenance procedures involve lots of swapping in and out of flight control components. Although multiple redundancy increases FCS integrity, it also increases to frequency of individual component failure due to increased component count. Procedures which minimize the mating cycles or the use of NPC connectors will be needed for fiber optics to survive normal aircraft maintenance.

3.12 Latency

Latency in a control system directly affects the phase margin of FCS and must be minimized.

Latency is the amount of time a signal takes to reach its destination once it has been sent. It is measured from the time the first bit of a message begins to be transmitted to the instant the last bit is received.

In legacy buses, the latency is dominated by the message timing. MIL-STD-1553 has a bit time of 1 µs and 429 has a bit time of 10 µs. Therefore, an ARINC-429 bus has between 1 ms and 81.6 ms depending on the message length. The 1553 bus has a required response time of 12 µs which

can be confused to mean a maximum latency. When the message timing is taken into account, the

latency of a 1553 bus is between 26 µs and 646 µs for a single, one way message. Round trip latency is up to twice this number, for a maximum of over 1 ms.

Modern buses have bit times in the 10’s of ns or less which practically eliminates transmission latency. Latency can also be introduced by the actual propagation time down the cable, buffering, or in the processing and signal conversion at each end of the cable. 1394 can have latencies of up to a few milliseconds. A modern ethernet protocol such as 10 GBase-T has a specified maximum latency of 1 µs. The conversion latencies and other protocol specified latencies are largely independent of

transmission media and do not favor fiber or copper.

30 Most of the latency in a control system comes from factors other than the data link. A processor in a modern FCC might operate at clock frequencies in the GHz, but a complicated control loop might take many clock cycles. Furthermore, the operating system running on such a computer must perform many tasks which compete for processor time. A very fast control loop on a FCC might update at a rate of up to 1 kHz or every 1 ms. This computational latency is larger than any transmission latency that might be present in a modern data link.

3.12.1 Transport Latency

Although it seems that light in a fiber would physically travel faster than electricity in a copper wire, this is not necessarily the case. An electrical signal traveling down a copper cable is an electromagnetic wave, just like light and travels a significant fraction of the speed of light. The speed of propagation in an electrical conductor is a function of the cable geometry and insulator type. Signals in ethernet cables travel between 40% and 70% of the speed of light. Typical optical

fiber has a refractive index of about 1.62 which corresponds to 1/1.62 or about 60% of the speed of light. In either case, a signal takes between 1.4 and 2.5 ns to travel a foot. Even for the largest aircraft, a single cable run of 100 ft would correspond to a transmission latency of a few µs in either

fiber or wire.

3.13 Redundancy

The flight control systems used in modern aircraft have become extremely complex to meet the requirements detailed above. Typically, there are two complete fault-tolerant flight control com- puter systems, a primary and an intentionally dissimilar backup[64]. These systems themselves are typically at least triple redundant and often quad redundant, which means there are typically six to eight independent flight control computers on any aircraft which are each capable of controlling the aircraft by themselves. Actuation systems are sometimes up to quad redundant as well, both

31 electrically and mechanically, depending on the effect of a control surface failure. Redundant and backup components are often intentionally of a different design than the primary components so that unknown design flaws are less likely to affect all redundant components at the same time. The need for dissimilar hardware was demonstrated by the first Ariane 5 rocket, which was lost because all its redundant hardware and software was identical. All the redundant systems failed simultaneously due to identical software bugs in each system.

This sort of overcapacity seems wasteful in aircraft, where weight and space are at a premium.

Structural margins in an aircraft are very small, typically 5-10% in order to minimize weight. How- ever, this redundancy is effectively mandated by the high integrity requirements of a flight control system. The redundancy levels in selected aircraft flight control systems are summarized in SAE

AIR 4235[56]. This document discusses the great lengths aircraft designers take to prevent a catas- trophic failure. The F-18 for instance, has between two and four position sensors present in each aircraft actuator. Quad-redundant sensors feed quad redundant computers over quad redundant links so that any three failures anywhere in the flight control computer system will result in an aircraft which is still able to fly.

3.14 Determinism

FCS must be deterministic. Determinism means that there is strictly limited uncertainty about what a flight control system will do in response to a given input. FAA and military standards and regulations require that FCS be mathematically proven to perform as designed. This requirement may be met in a variety of ways, but uncertainties must be strictly controlled.

32 3.15 Databus Protocols

Many of the characteristics necessary for a flight control system to function are controlled in the communications protocol used by its data bus. Flight control protocols must have high relia- bility, low latency, and high determinism. Examples of protocols currently used for flight control include, 1553, ARINC 429, and the newer AS 5643 (MIL 1394). These are very different protocols and each has its own unique characteristics. Optical replacements for these busies have been stud- ied may be appropriate ways to create a FBL control system. McDonnell Douglas and Honeywell demonstrated an optical 1553, 1773 in 1996 and a derivative, STANAG 3910, is used in the Eu- rofighter avionics[65, 66]. A newer optical bus, Avionics Full Duplex (AFDX) is used in the Airbus

A380 and Boeing 787 avionics systems and was actually designed for flight control systems. This newer bus could be a good choice for future flight control systems.

3.15.1 MIL-STD-1553

The 1553 bus is a 1 Mbit/sec multiplexed, data-bus. It is reverse-Manchester encoded with differential signaling. It is a command-response protocol which guarantees to the sender that the receiver has received the message. This bus has extremely high integrity due to its differential signaling and the Manchester encoding, which uses state transitions rather than voltage levels to signify individual bits. The actual bit rate is much lower than the raw bit rate would suggest. A large percentage of the bandwidth in a 1553 bus is taken up by the error checking and other protocol overhead. This means that in practice a single 1 Mbit/sec 1553 bus can only support a bandwidth of

200-300 kbit/sec[67].

33 3.15.2 ARINC 429

ARINC 429 is a unidirectional, one-to-many data-bus. It operates at a bit rate of 100 kbit/sec

This bus is popular in commercial aircraft. The single sender nature of this bus naturally precludes bus contention, but the bandwidth is extremely low. Two independent buses are required for two- way communication.

3.15.3 MIL-1394

The AS5643, or MIL 1394 bus is an improved version of IEEE 1394b which adds features necessary for flight controls[68, 69]. Features added on top of the 1394 spec include a Start of Frame

(STOF) packet for synchronization, static channel numbers, pre-assigned bandwidth, a Vertical

Parity Checking (VPC), and redundancy. All flexibility is intentionally removed from the bus for the sake of increased determinism.

AS 5643 utilizes a fixed frame rate which is synchronized based on the STOF packet. Unlike standard 1394, there is no resource manager to arbitrate between devices. Instead, bandwidth is preallocated with each device transmitting on an assigned channel with an offset and duration given in 1 µs increments. Triplex redundancy is explicitly used in MIL 1394.

VPC differs from standard parity checking in that it compares identical bits of successive words, rather than looking for bit errors in a single word. The VPC is computed by performing an XOR on each successive bit in a given position within the word. The final value is then bitwise negated and added to the end of the payload. This parity word can then be checked by the receiver after the CRC to verify that a CRC collision has not occurred. In this way, signal integrity of an entire message can be guaranteed above the 10−7 level provided by a standard 32-bit CRC.

34 CHAPTER IV

HISTORY OF OPTICAL FIBER IN AIRCRAFT

Fiber optics have been used in the mission systems of aircraft for many years. These links were introduced first in an experimental capacity[29]. Fiber optics are now commonly used in modern aircraft, spacecraft, and satellites and have begun to replace copper for data links for new equipment.

4.1 McDonnell Douglas AV-8B

In 1981, the US AV-8B Harrier, shown in Figure 4.1[70] was fitted with a single 43 ft fiber optic link as the primary data link for Communication, Navigation and Identification (CNI) with an elec- trical backup[29]. This was intended to be an experimental link with a short service life. However, the link had no operational failures in over 45,000 hours of flight time. A spare fiber was included in case of a field failure, however this spare was eventually removed because it was never used. The routing of the fiber was intentionally placed near the skin of the aircraft to facilitate replacement of the fiber. This placement exposed the fiber during certain routine maintenance activities and led to a large number of mechanical failures during aircraft maintenance. The routing was changed in a later version and no further mechanical failures were experienced. The link was finally removed during a scheduled refit in 1986 after 5 years of service.

35 This fiber optic link in the AV-8B was split in two places for manufacturing reasons and thus had three fiber cables and six fiber connectors. The connectors were of an SMA-905 NPC type and used relatively large 100/140 µm multimode fiber. These connectors were expensive and time consuming to install and were the main reason the fiber link was removed from service. At that time, it was believed that the elimination of epoxy from the connector would reduce cost. This turned out not to be the case. Modern, automatic polishers would have eliminated the need for hand polishing and reduced connector cost considerably without sacrificing performance.

The transmitter used in the Harrier was an 820 nm GaAlAs LED operated with an optical power of about 1.1 mW and a forward current of about 100 mA. The receiver was a PIN photo-diode with a responsivity of 0.3 A/w and a dark current of 0.05 nA at 30V bias. Both transmitter and receiver were sealed in hermetic TO-46 style cans. This transceiver was designed to have a reliability of only 10−3 failures per flight hour but extremely conservative estimates were used to arrive at that number. The actual failure rate was much lower, as no operational failures of either the fiber or transceiver were experienced in 5 years.

The fiber link in the harrier was a simple, low bandwidth optical data link. There were main- tenance problems initially due to poor routing, but the link performed well in actual operation.

This early example demonstrated that fiber optics could be used reliably in a harsh fighter aircraft environment.

Figure 4.1: AV-8B Harrier

36 4.2 General Dynamics F-16

The F-16 Block 60 is the latest upgrade to the Falcon, but is currently only deployed by the

United Arab Emirates, who funded its development. This upgrade has a completely new avionics package produced by Northrop Grumman, and includes a fiber optic avionic data bus[71]. The bus used is reported to be 1773 which would be a logical replacement to the standard 1553 bus because the protocol is the same and it is backward compatible in the case of older hardware which might be used. The F-16 was modeled in Section 6.2 and is pictured in Figure 4.2[72].

Figure 4.2: General Dynamics F-16 Fighting Falcon

4.3 Boeing FA-18

The current build of the F-18, the F/A-18E/F Super Hornet pictured in Figure 4.3[73], makes extensive use of fiber optic interconnects. At least three separate avionics systems are reported to use fiber channel links. According to the Harris Corporation, the F/A-18 E/F has two Fiber Channel

Network Switches manufactured with a total of 16 channels[74, 75]. The Tactical Aircraft Moving

37 Map Capability (TAMMAC), Active Electronically Steerable Scanned Array radar (AESA) and the

Weapons Replaceable Assembly (WRA) all have fiber channel links[74, 76, 77].

Figure 4.3: Boeing FA-18 Super Hornet

4.4 Lockheed Martin F-22

The F-22, pictured in Figure 4.4[78] first entered service in 2005. It boasts an extensive network of fiber optics produced by the Harris corporation. Its proprietary fiber high speed data bus (HSDB) and fiber optic transmitter/receiver components are used for both processor to processor and point to point communications[31]. Specific information about the avionics on this aircraft are difficult to obtain, but some details about the HSDB were published in 1991, just as the F-22 was being developed[79]. It is a 50 Mbit/s TDM bus and the paper even specifically suggests this might be an appropriate bus for FBL. Anecdotal evidence suggests there are as many as 160 individual fibers in the F-22.

38 Figure 4.4: Lockheed Martin F-22

4.5 Lockheed Martin F-35

The F-35 Joint Strike Fighter (JSF), Figure 4.5[80], is currently in development by Lockheed

Martin. It reportedly has an avionics bus with over 300 fibers. At least some of the bus is reported to be 2 GBit Fiber Channel[32]. It is also reported to have an electro-optical targeting system which communicates with the CPU via high speed fiber optics[81].

Figure 4.5: Lockheed Martin F-35

39 4.6 Eurofighter Typhoon

The Eurofighter Typhoon was first released in 2004, and is equipped with a fiber optic data bus called Stanag 3910. This bus is used for all mission critical systems and has flown for over 100,000

flight hours af of January 2011[26, 30]. This databus utlizes a TDM multiplexing scheme and a data rate of either 1 Mb/s or 20 Mb/s. The Eurofighter Typhoon is pictured in Figure 4.6[82].

Figure 4.6: Eurofighter Typhoon

4.7 Lockheed Martin EC 130 H Block 30

EC-130H Block 30, Figure 4.7[83], contains extensive fiber optics. This avionics suite was reportedly being rolled out to all USAF EC-130’s by the end of FY11. Reportedly, this aircraft contains more fiber optic terminations than any other production aircraft[84, 85].

40 Figure 4.7: Lockheed EC-130

4.8 Boeing 777

The Boeing 777, Figure 4.8[86], has an ARINC 636 fiber optic data bus. The ARINC 636 bus is a variant of the FDDI but has some differences[20]. The FDDI spec calls for a data rate of

125 Mbps with full duplex and a token ring topology. The FDDI spec also calls for the use of the

1270-1380 nm wavelength with LED sources and 62.5/125 nm graded index fiber or 8.7/125nm single mode fiber with laser sources. With either source, PIN diodes are used for receivers. The

LED source can transmit up to 2 km, and the laser source up to 40 km. ARINC 636 is using 62.5 um fiber probably because this is common in aircraft applications and the length of a fiber run in an aircraft will never exceed 2 km. Moreover, the larger fiber core is easier to maintain due to the lower tolerances required. The 636 bus uses NPC connectors for harsh locations and its optical power budget is designed to accommodate the higher resulting link losses.

41 Figure 4.8: Boeing 777

4.9 Boeing 787

The Boeing 787, Figure 4.9[87], has a fiber optic backbone, containing 110 individual fiber links and 1.7 km of fiber optic cable[88]. Fiber feeds the flight deck displays, Common Computing

Resources(CCR), and data concentrators. The 787 does not use the same physical protocol as the

777. Instead, it uses ARINC 664, which is a variant of the AFDX protocol designed for the Airbus

A380[89]. This standard uses a star topology with full duplex switched ethernet.

Figure 4.9: Boeing 787

42 CHAPTER V

EXPERIMENTAL FLY BY LIGHT

5.1 Nasa F-18

NASA modified an F-18 to use fiber optic command for a primary flight control surface in

1997[62, 90]. They replaced a single aileron with an optically commanded hydraulic actuator.

The resulting SMART actuator was designed by HR Textron, and was a standard servo-hydraulic actuator with an integrated servo-valve. The command link featured a dual redundant link module with fail-op/fail safe failure modes. Two SMA 905 connectors were attached to each link module, one for transmit and one for receive. The SMA 905 connector is a NPC connector, typically used with multimode fiber and possibly chosen due to its successful use in the AV-8B. The AV-8B used

100/140 µm multimode fiber in its cables.

The project experienced problems with the optical links during environmental and ground test- ing. The cable jackets shrunk during temperature cycling, and the fiber was crimped rather than epoxied into the connector ferrule. This low-cost, experimental assembly method allowed the fiber to protrude (piston) from the end of the connector. Also, the transmitter and receiver sensitivities were highly variable and had to be individually calibrated to prevent receiver saturation. This vari- ability appears to have been typical at the time[91, 92]. It should be noted that the AV-8B’s cables

43 underwent rigorous qualification testing and did not experience operational failures due to these problems in over 5 years of flight operations[29].

5.2 Gulfstream 550

An experimental Gulfstream 550, Shown in Figure 5.1[93], used a fiber optic transceiver in a direct replacement for the Fly by Wire controls to a spoiler. This transceiver was manufactured by

Defense Photonics and flew for in the experimental aircraft for approximately one year[94]. Defense

Photonics has since gone out of business, but at least two other companies are looking into similar technologies[95, 96].

Figure 5.1: Gulfstream 550

5.3 EC 135 Eurocopter

An experimental EC135 Eurocopter built in 2002 and shown in Figure 5.2[97], utilized an tra- ditional FBW control system with optical transmission media, possibly STANAG 3910. This was part of an advanced research vehicle and included the ability to simulate the flight characteristics of other helicopters, which was enabled by the high bandwidth of the optical transmission media. This system featured a quad redundant flight control computer to meet a 10−9 probability of failure[98].

44 Figure 5.2: Eurocopter EC-135

45 CHAPTER VI

RESEARCH WORK

6.1 Avionics Full Duplex

One obvious way to implement FBL is to use an optical bus to carry digital FBW signals. AFDX is an existing avionics protocol which utilizes cots fiber optic ethernet devices with modifications for a flight control environment. It was specifically designed for real time control systems. Airbus has expressed interest in using AFDX for flight controls and at least one other paper has explored this use of AFDX[99]. The research work described herein studies the use of AFDX in flight control systems(FCS).

Section 6.1 compares the AFDX bus to other avionics buses and lists a number of possible shortfalls of fiber optics for use in FCS. This section was published in a short paper by the author at the IEEE AVFOP conference in 2012[100]. The AVFOP paper contains recommendations for further research on AFDX and sections 6.2 - 6.3 are intended to address one such recommendation.

6.1.1 Fiber Optic AFDX versus Legacy Avionics Busses

A given fiber optic link weighs less than a copper cable even when substituted on a 1 for 1 basis[10]. However, the high reliability, electromagnetic interference (EMI) immunity and high

46 bandwidth of fiber links may offer several intriguing possibilities for modernization, such as reduc- tion in physical redundancy requirements through the use of Cross Channel Data Links (CCDL) or other advanced techniques. Moreover, the use of AFDX enables performance improvements due to increased bandwidth and/or word size; possibly resulting in higher loop closure rates, more precise sensor measurements, or additional bandwidth-limited capabilities[101].

AFDX has a raw bit rate of 100 Mbits/s vs. 1 Mbits/s for 1553 and 100 kbits/s for ARINC

429. The switched nature of AFDX naturally leads to contention when two or more devices trans- mit simultaneously. Although a properly designed AFDX bus will not drop packets, the required network buffers can and do lead to delays. One paper claims that an AFDX bus needs to be run at

10-20% capacity to minimize latency and jitter[102, 103, 104]. However, other buses suffer from similar problems. An ARINC 429 bus is single direction; two independent links are required for two-way communication between devices. Thus a single, realistically loaded AFDX network can transmit 100-200 times more than an ARINC 429 link. MIL-STD-1553 allows for two way commu- nication without contention using Time Division Multiplexing (TDM), but at the cost of extremely high overhead. Per the specification, the 1553 TDM scheme works best below 400 Hz and only marginally up to 3 kHz.

6.1.2 AFDX versus IEEE 1394

The Joint Strike Fighter, is known to use MIL 1394, which is comparable in performance to

AFDX[69]. Based on the written standards, AFDX and MIL 1394 are similar in bit rate and in- tegrity. MIL 1394 is slightly better in timing accuracy as shown in Table 6.1. The AFDX specifica- tion contains network examples in which the latency and jitter have maxima of 75 µs and 500 µs.

Better absolute limits are possible dependent on both hardware and network topology. MIL 1394 is a serial bus and out of order packets are not possible. AFDX can also guarantee ordinal packet flow through proper network design. Contention on a 1394 bus leads to an arbitration sequence which

47 lasts a fixed 3.7 µs. Repeated collisions can occur and lead to a bus reset after a fixed number of occurrences. AFDX does not have collisions, and contention is handled by buffering packets. 1394 and AFDX use the same Cyclic Redundancy Check (CRC) algorithm for error checking. They may provide differing levels of integrity based on packet sizes as well as assumed bit error rates due to physical construction. Signal integrity of 1394 versus AFDX is an area for future investigation. MIL

1394 uses an additional vertical parity check (VPC) to achieve the necessary integrity as described in Section 3.15.3.

Table 6.1: Comparison of AFDX Timing Parameters vs. Other Busses

Databus MaxLatency Max Jitter Contention Multiplex Type AFDX 75 µs* 500 µs* Yes Packet Switch MIL 1394 125 µs 25 ns* Yes TDM MIL 1553 646 µs** 8 µs No TDM ARINC 429 81.6 ms** 816 µs** No None * Dependent on network architecture ** Dominated by message timing (length x bandwidth)

6.1.3 AFDX Further Work

In-depth analysis on overall flight control system architecture is needed to assess data bus la- tency effects on overall system latency. FAA AC 25.1309 requires 10−9 likelihood of catastrophic failure for a safety critical flight control system. The Bit Error Rate (BER) and algorithm used to detect bit errors of on AFDX link do not provide this level of integrity, nor does it need to by itself.

Additional integrity layers need to be investigated.

Signal latency and jitter reduce phase and gain margins, time available for computation, and possibly sensor accuracy. These effects must be studied.

48 Non-Physical Contact (NPC) connectors were used on the ARINC 636 bus in the Boeing 777 to mitigate mating durability and cleanliness problems. ARINC 636 link budget requirements also address receiver saturation issues encountered in NASA/AFRL Smart actuator testing[105]. The use of NPC connectors needs to be studied to determine if this technology can bring similar benefits to AFDX.

6.1.4 AFDX Conclusion

AFDX was designed to enable the use of ethernet protocols for real time control systems. This testing will explore the use of AFDX in a flight control system. Several challenges specific to the

flight control system have been identified which require further analysis and testing. Future work should identify specific solutions and technical limitations which will guide the use of AFDX in this application.

6.2 Latency and Jitter Effect on Control Systems

Research was conducted studying the effect of latency and jitter on control systems. AFDX jitter is small compared to legacy buses, but still may have a measurable effect on high speed control loops. This effect was studied for two different control system models.

The first model investigated was a JSBsim simulation of an F-16 aircraft running on autopilot.

The F-16 was chosen because a mature aircraft model was included in the JSBsim package. The autopilot had to be written for this study and was adapted from one included for another aircraft.

JSBsim is an open source flight simulator developed by a retired NASA engineer and it is the default

flight dynamics model for the Flight Gear flight simulator. It can be run in a scripted mode and has an extremely flexible configuration system. Any property used by the model can be input, output, or modified in a variety of ways. This system allows the introduction of delays on any given sensor

49 input. This allowed for the study of varying amounts of latency on the flight control system and measurement of its effect on flight parameters.

A second system studied was a model of an induction heater system and PID control loop.

The heater model was previously used in the development of an actual control system and closely approximates the mechanical system modeled. It was validated against the physical system and the same PID controller studied is used currently to control the actual heater system. This is an appropriate model to use for a variety of reasons. Heaters are stable systems, and remove additional sources of noise which might be present in other types of control systems. The time scale of an induction heater is similar to a fast flight control system unlike most heater systems. Moreover, the system in question was already designed to communicate asynchronously over UDP and facilitated the introduction of arbitrary latency and jitter.

6.2.1 SAE AS 94900 FCS Requirements

The F-16 was chosen as a model to study because it is a common aircraft which is dynamically unstable and has readily available models for the aircraft and flight control system. The following is a simplified schematic representing the F-16 FCS, Figure 6.1. The figure includes inputs used to measure open loop and closed loop responses, as well as the aircraft control surface command in red. For both the pitch and yaw channels, there are two feedback signals. In either case there is an angular velocity feedback and a force feedback. The angular velocity channel is treated as the output for the bode plots generated.

There are a number of quantitative requirements for FCS performance included in the SAE AS

94900 standard. These requirements include single actuator gain and phase margins, deviation and overshoot limits during maneuvers, and deviation limits during straight and level flight.

50 Figure 6.1: F-16 Flight Control System Model

6.2.2 F-16 Control System Bode Plots

Bode Plots for all three axes are shown in Figures 6.7, 6.8, and 6.9. Both open loop and closed loop response were measured. Both yaw and roll axes are stable over the frequency span measured, but the pitch axis is highly unstable due to a resonant peak at about 11rad/sec or 1.75 Hz. All three axes are stable in the closed loop system.

51 Bode Blot Generation Method

Open loop response was measured by injecting a signal after the control system and comparing the input to the output. The closed loop resopnse was measured by injecting the same signal into the input to the control system instead. The injection points are shown in Figure 6.1. The output was taken to be the angular velocity of each channel.

Bode plots can be extracted from a model, a single frequency at a time by plotting the command and feedback and adjusting the output multiplier and time shift until the traces match. A representa- tive plot of the roll command and feedback signals is shown in Figure 6.2 It is sometimes easier to extract these constants from a parametric plot of input versus output, shown in Figure 6.3. The area inside the loop can be reduced by delaying the command signal or advancing the feedback signal.

This process could be automated, but for this research, the time shift was found by trial and error. A plot of the baseline signal with 119 ms of time shift is shown in Figure 6.4. The system gain can be determined by adding a multiplier to the command signal. A plot of the time shifted baseline with a multiplier of 0.276 applied to the command is shown in Figure 6.5. Note there is no area between the curve and that the slope matches the dotted reference line. The final plot versus time is shown in Figure 6.6.

Once the appropriate time shift and scale factor are determined, the system gain and phase shift can be extracted for a particular frequency. For the baseline plot shown, the excitation frequency is

11 Hz. The gain at 11 Hz is therefore, 20 ∗ log(0.276) = 11.18dB. A sine wave travels through 360 degrees per cycle, which means 360 deg ∗11Hz = 3960 deg /second. The phase shift is therefore

3960 deg /sec ∗ .119sec = 471 deg. This method of generating Bode plots has the benefit of working on real systems for which no model exists.

52 Figure 6.2: Representative Plot of Baseline Command and Feedback

Figure 6.3: Parametric Plot of Baseline Command and Feedback

53 Figure 6.4: Parametric Plot of Time Shifted Command and Feedback

Figure 6.5: Parametric Plot of Time Shifted and Scaled Command and Feedback

54 Figure 6.6: Final Time Shifted and Scaled Command and Feedback

Bode Plot Results

Figure 6.7: F-16 Pitch Axis Open Loop Diagram

55 Figure 6.8: F-16 Yaw Axis Open Loop Diagram

Figure 6.9: F-16 Roll Axis Open Loop Diagram

56 The F-16 is unstable in two and possiby three axes. The yaw and roll axes are measurably unstable. Gain and phase margin could not be assessed for the pitch axis because the gain did not cross unity within the measurable frequency range but was always positive. Likewise, the pitch axis phase lag did not cross 180°.

The yaw axis had a negative phase margin of -92° at 10.5rad/sec The required phase margin is +45° and the limit of stability is 0dB. The gain margin was also negative, about -10 dB at about

5 Hz. The required gain margin is +6 dB with 0 dB being the limit of stability. The roll axis is unstable as well. The gain margin is about -2 dB and the phase is -34°.

6.2.3 F-16 Level Flight Handling

In steady flight there are flight smoothness requirements. The aircraft must not produce more than 0.05 g vertical or 0.02 g lateral acceleration. It must not sideslip more than 1° or have more than 0.5° of pitch angle. It must control altitude to within 30 ft, heading to within 0.5° and speed within 0.01 mach or 2% of indicated speed whichever is greater.

The aircraft meets these specs in level flight. A selection of the requirements most adversely affected by lag are shown in Figures 6.10 and 6.11. Both figures have red boxes showing the allowable limits, including a three second settling time. The absolute limit for latency in the FCS is

52 ms in level flight. At this level, the acceleration in both axes is approaching the allowable limit, as shown in Figure 6.10. Altitude and velocity are also shown, but they are not noticably affected.

The plane goes unstable with 53 ms systemwide latency. The accelerations in both axes both ring and steadily increase in magnitude as seen in Figure 6.11. Eventually, this instability shows up in the velocity and altitude traces as well.

57 Figure 6.10: F-16 Level Flight Latency Limit

58 Figure 6.11: F-16 Level Flight Unstable

59 6.2.4 F-16 Heading Change Handling

During a heading change, the maximum deviation is 1° with a 2.5° overshoot. The roll rate must be less than 5°/sec. During a bank, the sideslip angle must be less than 2° and the maximum lateral g force is 0.03 g. While rolling, the maximum lateral acceleration is 0.05 g. The maximum pitch deviation is 1° and the settling time must be less than 3 seconds. During a heading change, the speed must not change more than 10 knots or 2% of indicated speed, whichever is greater.

The requirements for flight handling during a heading change are mostly met by the baseline system. However, the accelerations seen are greater than allowed by the specification even with no added latency as seen in Figure 6.12. It is likely the spec was not written with fighter jets in mind.

For this reason, the lateral acceleration limit was not used as a measure of aircraft instability for the heading change.

The latency limit for a heading change was 49 ms as seen in Figure 6.13. In this plot, the heading deviation, roll rate, and velocity all remain within the specified limits. The aircraft becomes unstable with this maneuver at 50 ms of systemwide latency, Figure 6.14. The heading and roll rate rapidly exceed the limits, though the velocity does remain within spec for the entire simulation. This is a slightly lower limit than for level flight.

60 Figure 6.12: F-16 Baseline Accel

61 Figure 6.13: F-16 Heading Change Latency Limit

62 Figure 6.14: F-16 Heading Change Unstable

63 6.2.5 F-16 4 ft/s Altitude Change Handling

During an altitude change there is a maximum of 0.5 g normal, and the speed must not change more than 10 knots or 2% of indicated speed. At a ramp rate of 4 feet per second, there must be less than 20% overshoot and less than 175% pitch overshoot.

Figure 6.15: F-16 4 ft/s Altitude Change Latency Limit

64 Figure 6.16: F-16 4 ft/s Altitude Change Unstable

6.2.6 F-16 30 ft/s Altitude Change Handling

During a 30 foot per second ramp, the deviation must be less than 5° per second from the commanded pitch.

65 Figure 6.17: F-16 30 ft/s Altitude Change Latency Limit

Figure 6.18: F-16 30 ft/s Altitude Change Unstable

66 6.2.7 F-16 FCS Latency Conclusion

The F-16 was measured to be dynamically unstable in the pitch axis in the open loop condition.

The FCS brings this under control and the aircraft is stable with the FCS in the loop. Introducing latency could be expected to loosen the coupling between the FCS and the aircraft and thereby destabilize the aircraft, and this affect is demonstrated.

The pitch axis goes unstable with 53 ms of introduced latency in level flight. One maneuver, the heading change, reduces this instability threshold to 50 ms. Other maneuvers do not cause instability at a lower latency. 50 ms is vastly greater than the latencies introduced by a properly designed AFDX bus. The AFDX bus will therefore not affect system stability due to transport latencies.

6.3 Induction Heater Model and PID Controller

6.3.1 Heater Model and Validation

As mentioned, the heater model studied was an actual, validated physical model of an induction heater system. An induction heater approximates a single control loop in a FCS in a variety of ways, including a short time constant, and rapid, catastrophic failure modes. The model used was designed to closely match the physical response of the real physical system. The resulting PID loop and associated interface is currently in use, controlling the heating for a series of high temperature tensile and fatigue tests as shown in Figure 6.19.

67 Figure 6.19: Induction Heater Under Closed Loop Control

6.3.2 Physical Model Details

The heater block was physically modeled as a lumped system using estimated energy flow, masses, and capacities for the components used. The model was updated every 2 ms. A diagram of the heater can be found in Figure 6.20 and of the heater system in Figure 6.21 The LabVIEW code for updating the command, introducing arbitrary delays without disturbing the physical model is shown in Figure 6.22. The model is run in real time and the command signal is passed asyn- chronously via UDP. The system updates its state every 2 ms whether or not the command signal changes and updates the process variable which is itself passed asynchronously back to the con- troller. In this way, the command and feedback are decoupled from either the controller or the system. Arbitrary latencies and jitter can be introduced into the control loop and the effect of these varying delays can be studied.

68 Figure 6.20: Induction Heater Physical Model

Figure 6.21: Induction Heater System Physical Model

69 Figure 6.22: Physical Model and Jitter and Latency Introduction

70 The LabVIEW GUI is shown in Figure 6.23. The program sends a sinusoidal command signal to the heater model as described and reads back the process variable on a second channel. There is a delay between command and feedback signal due to a fixed sampling period and the command signal is shifted to remove this roughly constant delay. The program was designed to collect data for

100 seconds at a time and automatically calculate error values between the command and feedback signal. Latency was introduced on the system side by adding a fixed time delay before sending the feedback signal to the controller. Jitter was created by multiplying white noise with a magnitude between 0 and 1 by the peak delay value chosen and again delaying the signal by this amount. The artificially introduced delay was then the sum of this latency and jitter. Data were collected while varying latency only, peak jitter only, and equal latency and peak jitter.

Figure 6.23: LabVIEW Error Measurement GUI

71 6.3.3 Latency and Jitter Correlation Results

The resulting data was studied to determine the relationship between delay and error. For this system, with this noise type, the resulting system error is almost directly proportional to the RMS magnitude of the introduced delay as shown in Figure 6.24. The RMS value for introduced latency is just the constant latency value used for each run, while the RMS magnitude of the latency is less than the peak latency. This makes good sense in hindsight because the RMS captures the average magnitude of a time varying signal. It also means that the effect of jitter can be predicted for this type of bounded noise. Deterministic ethernet has fixed latencies and bounded jitter similar to the system used. Likely, jitter would have a much more sparse distribution, occurring only when com- munication system was under heavy load. However, this continuous, severe system jitter enabled the measurement of jitter in a worst-case case situation.

The correlation between position error and latency could be used to predict the effect of a given latency and jitter on a control system. This effect could be modeled as sensor uncertainty and should be compatible with standard modeling procedures. This relationship between error, latency, and jitter is only valid for the specific command frequency, but slower frequency inputs should have correspondingly smaller positional errors.

72 Figure 6.24: Best Fit of Latency and Jitter On PID Control System

6.4 Research Summary

This research shows that AFDX would be an order of magnitude bandwidth performance im- provement over legacy avionics buses. It is comparable in performance to another avionics bus,

MIL 1394, known to be used in the JSF. While much work would need to be done to build an actual

AFDX based FCS, this work lays some of the groundwork for such a design.

Research summarized in this paper studies the effect of latency on the F-16 FCS. Latencies required to destabilize this aircraft are not possible for the flight control system studied and can therefore be ignored. The software used to model this aircraft was not capable of introducing timing jitter so only latency was studied for these systems.

73 Further research attempted to measure the effect of latency versus jitter on a realistic PID loop controlled system. This research shows that the errors introduced by latency and jitter are propor- tional to the RMS magnitude of these values. This means that the effect of jitter was much smaller than latency of equivalent magnitude. The flight control system studied was tolerant of latencies far in excess of the 500 µs, single packet latency possible in an AFDX bus. This in turn implies that the flight control system studied in the previous sections could easily tolerate both the maximum latency and jitter introduced by an AFDX bus.

74 CHAPTER VII

CONCLUSION

Fiber optics have become a robust, mature technology and are currently the defacto standard for high performance communications. Many aircraft now contain fiber optics, and experience with this media in an aircraft environment is becoming more and more common. Digital flight control systems have matured to the point where nearly all modern aircraft have digital and even Fly by

Wire controls. There has been experimentation with Fly by Light over the years, and it is time for a production Fly by Light aircraft to realize the unique benefits of this media in an aircraft flight control system.

A number of recommendations can be made based on the literature survey and research con- ducted in this paper:

1) Wavelength division multiplexing should be implemented to maximize the bandwidth and channel separation between differing systems in the aircraft. WDM could allow devices of mixed criticality, eg. flight-critical, mission-critical, non-critical, to share the same fiber without interfer- ence.

2) 62.5 nm core multimode fiber has proven to be reliable in multiple aircraft and the relatively lower bandwidth and higher losses would not have a negative impact on flight control systems.

75 Moreover, multimode fiber is far easier to maintain - a primary concern in modern aircraft. Mul- timode fiber occupies the same physical space as singlemode fiber and dispersion should not be a problem in the short distance available within an aircraft.

3) Physical contact connectors cannot survive current routine aircraft maintenance practices.

Non physical contact and possibly expanded beam connectors should be used in aircraft flight con- trol systems. This will allow fiber links to survive the necessary, frequent mating cycles which take place during flight control system maintenance. As an alternative, systems could be designed in which flight modules could be removed without demating the fiber optic interconnects.

4) Military and commercial standards need to be developed for fiber optics in aircraft in gen- eral, and flight critical systems in particular. The lack of standards has been documented and will continue to limit fiber use in aircraft.

5) AFDX is one appropriate interconnect media to implement Fly by Light. Additional error detection such as a vertical parity check, used in MIL 1394 and other high reliability systems, would provide the 10−9 integrity necessary for flight control. The research conducted in this report shows that the small, bounded, non-deterministic latency and timing jitter present in an AFDX bus would not adversely affect a flight control system.

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