Parent Child Unmanned Aerial Vehicles and the Structural ...

Parent Child Unmanned Aerial Vehicles and the Structural

Dynamics of an Outboard Horizontal Stabilizer Aircraft

by

Jason Kepler

Submitted to the Department of Aeronautics and Astronautics inpartial fulfillment of the requirements for the degree of

Master of Science in Aeronautics and Astronautics

at the

MASSACHUSETTS INSTITUTE OF TECHNOLOGY

June 2002

© Massachusetts Institute of Technology. All Rights Reserved.

A u th o r ....................................................... .. F- ----- --- -----Departni n autics an A ronautics

A May 12, 2002

Certified by ....................................f &hn J. Deyst

Professor of Aeronautics and AstronauticsThesis Supervisor

A ccepted by .............................Wallace E. Vander Velde

Professor of Aeronautics and AstronauticsChairman, Committee of Graduate Studies

MASSACHOSETTS WNSTITUTEOF TECHNOLOGY

AERO AUG 13 2002

LIBRARIES

Parent Child Unmanned Aerial Vehicles and theStructural Dynamics of an Outboard Horizontal

Stabilizer Aircraftby

Jason Kepler

Submitted to the Department of Aeronautics and Astronautics onMay 12, 2002, in partial fulfillment of the requirements for thedegree of Master of Science in Aeronautics and Astronautics

Abstract

In the fall of 1998, MIT and Draper Laboratory formed a partnership program called Par-ent Child Unmanned Aerial Vehicle (PCUAV) to provide a means of providing upclosesurveillance at a distance. The premise of the project was to create a tiered system of coop-erative autonomous aircraft. A large Parent aircraft was designed to carry a smaller Miniaircraft to a target site and release it to descend for upclose surveillance. Meanwhile, theParent provides a communications link between the Mini and a ground station at the pointof departure. At the completion of a surveillance mission the Parent retrieves the Mini andcarries it home.

This thesis discusses the system components for the PCUAV project, specifically concen-trating on the flight vehicles. The design, building, and flight testing phases for each vehi-cle are detailed. Special attention is given to the Parent vehicle, which utilizes anOutboard Horizontal Stabilizer (OHS) configuration. The structural dynamics and bothaeroelastic and servo aeroelastic properties of the plane were studied using Aswing andare reported on here.

Thesis Supervisor: John J. DeystTitle: Professor of Aeronautics and Astronautics

Thesis Supervisor: Marthinus C. van SchoorTitle: Lecturer, Department of Aeronautics and Astronautics

3

Table of Contents

Table of Contents ................................................................................................... 5List of Figures ....................................................................................................... 7List of Tables ............................................................................................................ 11Acknow ledgm ents .................................................................................................. 13List of Acronym s and Sym bols .............................................................................. 151. Introduction ....................................................................................................... 17

1.1 Background and M otivations of PCUAV .............................................. 171.2 Background of the Outboard Horizontal Stabilizer Configuration ..... 191.3 Thesis Overview ................................................................................... 20

2. PCU AV System ................................................................................................... 212.1 Chapter Overview ................................................................................. 212.2 The PCUAV System Concepts ................................................................ 21

2.2.1 Typical Flight .............................................................................. 212.2.2 Key Enablers .............................................................................. 232.2.3 Reintegration .............................................................................. 24

2.3 Com ponents of PCUAV ........................................................................ 272.3.1 Parent Vehicle ............................................................................ 282.3.2 M ini Vehicle ............................................................................... 292.3.3 Avionics Testbed Aircraft ............................................................ 312.3.4 Payload Delivery Vehicle ............................................................ 322.3.5 Mini-Parent Integration Mechanism (MPIM) ............... 332.3.6 M id-Air Recovery System .......................................................... 362.3.7 Com m unications and Surveillance .............................................. 372.3.8 Flight Avionics ............................................................................ 38

2.4 Chapter Sum m ary ................................................................................ 433. UAV Building and Testing .................................................................................. 45

3.1 Chapter Overview ................................................................................ 453.2 OH S Parent Vehicle .............................................................................. 45

3.2.1 Advantages and Disadvantages of the OHS ................................ 453.2.2 OH S Parent Design Process ....................................................... 473.2.3 OHS Construction ........................................ 483.2.4 OH S Testing and Updating ......................................................... 52

3.3 M ini Vehicle .......................................................................................... 563.3.1 M ini Design Process ................................................................... 563.3.2 M ini Construction ........................................................................ 573.3.3 M ini Testing ................................................................................. 60

3.4 Avionics Testbed Aircraft ...................................................................... 613.4.1 Advantages and Disadvantages of the ATA ................................ 613.4.2 W ork done on ATAs ................................................................... 62

3.5 ATA Testing ......................................................................................... 633.6 Chapter Sum m ary ................................................................................ 63

4. Structural M odeling of the Parent ..................................................................... 654.1 Chapter Overview ................................................................................. 65

5

4.2 Structural and Inertial Properties of the Parent ..................................... 654.2.1 Area M oments of Inertia ............................................................. 654.2.2 W eight and M ass M oment of Inertia ............................................ 70

4.3 Analysis Process .................................................................................. 714.4 Natural Frequencies and M ode Shapes of the Parent ........................... 754.5 Chapter Overview ................................................................................ 77

5. Aeroelasticity .................................................................................................... 795.1 Chapter Overview ................................................................................ 795.2 Aeroelasticity of the Parent UAV ......................................................... 795.3 Flight Dynamics of Parent ..................................................................... 825.4 Servo Aeroelasticity of Parent .............................................................. 845.5 Chapter Summary ................................................................................ 88

6. Summary and Conclusions ................................................................................. 896.1 Thesis Summary ................................................................................... 89

6.1.1 PCUAV System Summary ............................................................. 896.1.2 Suggestions for UAV Improvements ......................................... 906.1.3 Flight Tests .................................................................................. 926.1.4 Structural Modeling of OHS Aircraft Summary and Conclusions . 92

Appendix A Vehicle Drawings ......................................................................... 97A. 1 Three View Drawings of PCUAV Parent Aircraft ............................... 98A.2 Three View Drawings of PCUAV NGM I ............................................... 99A.3 Three View Drawings of PCUAV NGM II ............................................. 100A.4 Three View Drawings of PCUAV ATA I&II .................... 101A.5 Building Plans for NGM II ...................................................................... 102A.6 W ing M oment of Inertia Spreadsheet ...................................................... 104

Appendix B Aswing and Related Code for the Parent .......................................... 105B.1 Description of Aswing ............................................................................. 105B.2 Aswing Code for the Parent ..................................................................... 107

Appendix C Aswing Results ................................................................................. 113C.1 M ode Shapes ............................................................................................ 114C.2 Bode Plots ................................................................................................ 125

6

List of Figures

Figure 1.1 Multi-Tiered System Concept ......................................................... 18Figure 1.2 Outboard Horizontal Stabilizer Parent Vehicle ................................ 19Figure 2.1 Communications Hierarchy for PCUAV .......................................... 23Figure 2.2 Phase One of Reintegration .............................................................. 25Figure 2.3 Phase Two Detection and Navigation System ................................. 26Figure 2.4 Parent Aircraft Inside a Dodge Caravan .......................................... 29Figure 2.5 (Left) NGMI, (Right) NGMII .......................................................... 30Figure 2.6 The First Avionics Testbed Aircraft ................................................. 32Figure 2.7 Payload Delivery Vehicle ................................................................. 33Figure 2.8 Original M PIM Design ........................................................................ 34Figure 2.9 Parent Aircraft with MPIM Attached .................................................. 34Figure 2.10 D etail of M PIM ................................................................................ 35Figure 2.11 MARS Directional Finder ................................................................ 37Figure 2.12 Rover with Surveillance Equipment on Top .................................... 38Figure 2.13 NGMII Flight Control Avionics ....................................................... 39Figure 2.14 Parent's Avionics Structure .............................................................. 43Figure 3.1 Vortex Induced Angle of Attack at Tail Position ............................ 46Figure 3.2 Parent's Spar D etail .......................................................................... 49

Figure 3.3 Parent Wing Composite Layup ....................................................... 50Figure 3.4 (Left) Author with Parent's Tail, (Right) Parent's Fuselage Frame .... 51Figure 3.5 Bending Moment in Parent's Wing ................................................ 53Figure 3.6 Second Landing of OHS Parent ....................................................... 54

Figure 3.7 Avionics Inside NGMII Fuselage ..................................................... 58Figure 3.8 Cross Section of NGMII Wing ....................................................... 59Figure 4.1 Cross Section of the Parent's Tail Booms ....................................... 66Figure 4.2 Cross Section of the Parent's Wing ................................................. 68Figure 4.3 Aswing Geometry for Parent ............................................................ 72Figure 4.4 Velocity Sweep of Parent ................................................................ 74

Figure 4.5 Root Locus Plot for Parent .............................................................. 75Figure 5.1 First flutter mode of OHS Parent ..................................................... 80Figure 5.2 Root Locus Plot of OHS Parent with Three Pound Weights

on Each Tail and Counterweight Attached to Fuselage ................... 81Figure 5.3 Root Locus Plot of OHS Parent with Three Pound Weights

on Each Tail and Counterweight Attached to Wingtips' Leading Edges 82Figure 5.4 Blow Up of Root Locus Near Origin .............................................. 83Figure 5.5 Bode Plots of Parent Roll Rate Response to Unit Aileron Input,

A irspeed = 70 ft/sec., @ S.L. .............................................................. 85Figure 5.6 Bode Plots of Parent Pitch Rate Response to Unit Elevator Input,

A irspeed = 70 ft/sec., @ S.L. .............................................................. 86Figure 5.7 Bode Plots of Parent Yaw Rate Response to Unit Rudder Input,

A irspeed = 70 ft/sec., @ S.L. .............................................................. 87Figure A. 1 Orthogonal Views of OHS Parent ................................................... 98Figure A.2 Orthogonal Views of New Generation Mini .................................... 99

7

FigureFigureFigure

A.3A.4A.5

Figure A.6Figure C.1

Figure C.2

Figure C.3

Figure C.4

Figure C.5

Figure C.6

Figure C.7

Figure C.8

Figure C.9

Figure C.10

Figure C.11

Figure C.12

Figure C.13

Figure C.14

Figure C.15

Figure C.16

Figure C.17

Figure C.18

Figure C.19

Figure C.20

Figure C.21

Orthogonal Views of Second New Generation Mini ..........................Orthogonal Views of Two Avionics Testbed Aircraft ........................Building Plans for NGM II Fuselage ..............................................

100101102

Building Plans for NGM II W ing and Tail .......................................... 103First Mode Shape of OHS Parent, Asymmetric VerticalTail Boom Bending, Airspeed = 70 ft/sec, @ Sea Level .................... 114Second Mode Shape of OHS Parent, Asymmetric HorizontalTail Boom Bending, Airspeed = 70 ft/sec, @ Sea Level .................... 115Third Mode Shape of OHS Parent, Symmetric Wing Bending,Airspeed = 70 ft/sec, @ Sea Level ...................................................... 116Fourth Mode Shape of OHS Parent, Asymmetric Wing Twist,Airspeed = 70 ft/sec, @ Sea Level ...................................................... 117Fifth Mode Shape of OHS Parent, Symmetric Wing Twist,Airspeed = 70 ft/sec, @ Sea Level ...................................................... 118Sixth Mode Shape of OHS Parent, Asymmetric HorizontalStabilizer Bending, Airspeed = 70 ft/sec, @ Sea Level ...................... 119Seventh Mode Shape of OHS Parent, Symmetric HorizontalStabilizer Bending, Airspeed = 70 ft/sec, @ Sea Level ...................... 120Eighth Mode Shape of OHS Parent, Second Wing Bending,Airspeed = 70 ft/sec, @ Sea Level ...................................................... 121Ninth Mode Shape of OHS Parent, Fore-Aft Wing Bending,Airspeed = 70 ft/sec, @ Sea Level ...................................................... 122Tenth Mode Shape of OHS Parent, Asymmetric VerticalTail Bending, Airspeed = 70 ft/sec, @ Sea Level ............................... 123Eleventh Mode Shape of OHS Parent, Symmetric VerticalTail Bending, Airspeed = 70 ft/sec, @ Sea Level ............................... 124Gain Plot of Roll Rate vs. Aileron Input Frequency forFlexible OHS Parent, Airspeed = 70 ft/sec, @ Sea LevelPhase Plot of Roll Rate vs. Aileron Input Frequency ForFlexible OHS Parent, Airspeed = 70 ft/sec, @ Sea LevelGain Plot of Pitch Rate vs. Elevator Input Frequency,Flexible OHS Parent, Airspeed = 70 ft/sec, @ Sea LevelPhase Plot of Pitch Rate vs. Elevator Input Frequency,Flexible OHS Parent, Airspeed = 70 ft/sec, @ Sea LevelGain Plot of Yaw rate vs Rudder Input Frequency ForFlexible OHS Parent, Airspeed = 70 ft/sec, @ Sea LevelPhase Plot of Yaw Rate vs. Rudder Input Frequency ForFlexible OHS Parent, Airspeed = 70 ft/sec, @ Sea LevelGain Plot of Roll Rate vs. Aileron Input Frequency for

125

................... 126

................... 127

................... 128

................... 129

................... 130

Rigid OHS Parent, Airspeed = 70 ft/sec, @ Sea Level ....................... 131Phase Plot of Roll Rate vs. Aileron Input Frequency ForRigid OHS Parent, Airspeed = 70 ft/sec, @ Sea Level ....................... 132Gain Plot of Pitch Rate vs. Elevator Input Frequency,Rigid OHS Parent, Airspeed = 70 ft/sec, @ Sea Level ....................... 133Phase Plot of Pitch Rate vs. Elevator Input Frequency,Rigid OHS Parent, Airspeed = 70 ft/sec, @ Sea Level ....................... 134

8

Figure C.22 Gain Plot of Yaw rate vs. Rudder Input Frequency ForRigid OHS Parent, Airspeed = 70 ft/sec, @ Sea Level ....................... 135

Figure C.23 Phase Plot of Yaw Rate vs. Rudder Input Frequency ForRigid OHS Parent, Airspeed = 70 ft/sec, @ Sea Level ....................... 136

9

List of Tables

Table 3.1 OH S Final Dim ensions...................................................................... 48Table 3.2 Final Dimensions of NGM II .............................................................. 60Table 3.3 Comparison of ATAI and ATAII Dimensions................................... 62Table 4.1 Cross Sectional Flexural Properties of Parent ................................... 68Table 4.2 Weights and Mass Moments of Inertia of Parent Components...... 70Table 4.3 Mode Shapes and Natural Frequencies of Parent,

Airspeed = 70 ft/sec, @ S.L. ............................................................. 76Table 5.1 Computed Flight Dynamic Modes Compared with Aswing Results .... 84

11

Acknowledgments

First I would like to thank the Lord God for his help and guidance through this project.He has provided me with more opportunity than I could have hoped for. Without Him Iwould have floundered long ago.

I also would like to thank my grandfather, Dave Dolese, who passed away during thewriting of this thesis. He was a good friend and did a lot to encourage me in my academicpursuits. Without his financial support I would never have been able to attend as fine ofinstitutions as I did.

Thank you to Professor John Deyst for his support and guidance. You treated us likeyour own children, which was more than I ever expected from an advisor at MIT. Also,thank you Dr. Tienie van Schoor for your time and effort, I really learned a lot from youboth in class and while working on this thesis. Thanks to Professor Mark Drela for yourhelp and for creating the software I needed to finish this thesis; I hope I can continue todraw on your vast knowledge.

I would like to thank Don Weiner for his help in the shop, his experience was invalu-able, and his grandfatherly advise on life was dearly appreciated. Thanks to Dick Per-dichizzi for his help with the windtunnel and all of the other facilities. Thanks to Col.Young for his aircraft building advice and infinite supply of stories.

Thank you to the other members of the PCUAV team. Francois Urbain was a greatfriend who I enjoyed building airplanes with. Sanghyuk Park never ceases to amaze mewith his hard work and seemingly unlimited intelligence. Thomas Jones was always goodfor a funny story. I appreciated the discussions I had with Alexander Olmenchenko aboutRussia and hockey, with Sarah Saleh about England and mad cows, and Richard Pourtrelabout flying. Good luck to Richard on his new flying career. Thanks to Damien Jourdan,it was nice to have another brother in Christ during the last months of the project.

Thank you to my parents, Chris and Susan Kepler, and my sister Caity, who all tooktime to come from across the country to visit frequently, and were always interested inwhat I was doing here. My parents inspiration was vital to my drive to excel in life andengineering.

Finally, I would like to thank the most important person in my life, my wife Christy.Thank you for marrying me and moving across the country to the big city so that I couldfulfill my dream. You have been the most supportive and loving wife a man can ask for. Ilove you.

13

List of Acronyms and Symbols

Acronymsa.c.Aero/AstroATA(I)&(II)AVLc.g.CPUDGPSDOSFMGM-15GPSIMUJPEGMACMARSMAVMITMPIMNASANACANGM(I)&(II)OHSPCMPCUAVPDVR/C or RCRFRxSBCS.L.UAVWASPWLAN

Symbols

(0Ab

Aerodynamic centerDepartment of Aeronautics and AstronauticsAvionics Testbed Aircraft, (I)&(II) refer to the different versions builtAthena Vortex LatticeCenter of GravityComputer Processing UnitDifferential Global Positioning SystemDisk Operating SystemFrequency ModulationGilbert Morris AirfoilGlobal Positioning SystemInertial Measurement UnitJoint Photographic Experts GroupMean Aerodynamic ChordMid-Air Retrieval SystemMicro Autonomous VehicleMassachusetts Institute of TechnologyMini-Parent Integration MechanismNational Air and Space AdministrationNational Advisory Committee for AeronauticsNew Generation Mini, (I)&(II) refer to the different versions builtOutboard Horizontal StabilizerPulse Coded ModulationParent and Child Unmanned Aerial VehiclePayload Delivery VehicleRemote ControlRadio FrequencyReceiverSingle Board ComputerSea LevelUnmanned Air VehicleWide Area Surveillance ProjectileWireless Local Area Network

Linear DeflectionDamping RatioEigenvalueAngular DeflectionFrequencyAreaSide Length

15

b, Wing Span

CL Coefficient of Liftcw Wing ChordD Dragd DistanceE Young's Modulus of ElasticityF ForceG Modulus of Rigidity

g Acceleration of Gravity on Earthh Side LengthHz Hertz (cycles per second)Ixx Area Moment of Inertia About x-axisIYY Area Moment of Inertia About y-axis

Izz Area Moment of Inertia About z-axisL Length, LiftL p Roll Moment Due to Sideslip AngleLr Roll Moment Due to Yaw RateLht Distance from Wing Quarter Chord to Horizontal Tail Quarter ChordLvt Distance from Wing Quarter Chord to Vertical Tail Quarter ChordJO Polar Moment of InertiakHz Kilohertz (thousand cycles per second)m MassMCC Pitch Moment Due to Angle of AttackMa Pitch Moment Due to Change in Angle of AttackMq Pitch Moment Due to Pitch RateMHz Megahertz (million cycles per second)No Yaw Moment Due to Sideslip AngleNr Yaw Moment Due to Yaw Rater radius

Sht Vertical Tail AreaSvt Horizontal Tail AreaSW Wing AreaT Torqueuo Initial VelocityV Volt

Vht Horizontal Tail Volume CoefficientVvt Vertical Tail Volume CoefficientX Coordinate Along x-axisY Coordinate Along y-axisYp Sideways Acceleration Due to Sideslip Angle

Yr Sideways Velocity Due to Yaw RateZ Vertical Acceleration Due to Angle of Attack

16

Chapter

1Introduction

1.1 Background and Motivations of PCUAV

In the fall of 1998, Draper Laboratory and MIT formed a project under their partnership,

known as the Parent Child Unmanned Aerial Vehicle (PCUAV) project. This was the sec-

ond of such projects formed under the Draper/MIT partnership, the first being the Wide

Area Surveillance Projectile (WASP) project that has since been taken over exclusively by

Draper Laboratory.

In forming the PCUAV project, the members of the team were addressing what was

perceived as an important aspect of military surveillance, namely to observe some point of

interest, at close range, from a distance. Current long range UAV surveillance aircraft are

inadequate for getting right down in the thick of where things are happening. Alterna-

tively, smaller micro air vehicles, with high maneuverability and low detectabillity, have

had such a short range and endurance that they had to be launched at close proximity to

the point of interest.

The goal of PCUAV is to create a system that would provide the benefits of both large

UAVs and small Micro Air Vehicles (MAVs) without incurring their associated disadvan-

tages. To do this, a concept evolved to use large scale UAVs to transport smaller aircraft to

17

a point of interest, launch them, provide a communications relay back to a ground station,

and then retrieve and bring back the smaller vehicles. When this concept was expanded to

include vehicles of many different sizes, the solution turned into a tiered system in which a

large Parent vehicle with an extended range and endurance could transport and launch

smaller Child, or Mini, vehicles as well as even smaller MAVs. The surveillance of each

vehicle could then be communicated back through each of the tiers to an operator at the

point of departure. When the mission is complete, either all or some of the aircraft could

be retrieved by the Parent and brought back for reuse.

Tier 1

T ier 2 - - - - - - - - - - - - - - - - - - -

Tier 3 a

Figure 1.1 Multi-Tiered System Concept

To accomplish such a mission, many key technologies must be investigated and dem-

onstrated. Some of them are as follows:

- Autonomous navigation of aircraft of various sizes

- Rendezvousing and reintegrating autonomous vehicles

" Visual surveillance and transmission of images between multiple aircraft

During the first two years of the PCUAV project the team concentrated on designing

the aircraft for the system and developing and building the guidance and control systems

18 Chap~ter 1: Introduction

for those vehicles. In the third year, when the author joined the project, and continuing

into the fourth year, the aircraft were built and flown, the control systems were validated

and work progressed toward demonstrating autonomous docking of the Parent and Mini

vehicles. The other technology that has been investigated and demonstrated during the

length of the project is surveillance and transmission of images between ground cameras,

a flying UAV, and an operator station.

1.2 Background of the Outboard Horizontal Stabilizer Configuration

The outboard horizontal stabilizer configuration, or OHS, was chosen for the parent vehi-

cle for reasons described in Chapters 2 and 3. This configuration looks somewhat foreign

to a first time observer. The aircraft has a center fuselage with no tail attached to it. Tail

booms extend rearward from each wing tip to both vertical and horizontal surfaces. The

horizontal surface extends outboard of the tail boom, leaving empty space between the tail

booms. As might be expected, this configuration has some unusual aerodynamics associ-

ated with it. See Figure 1.2 for a picture of PCUAV's Parent vehicle.

Figure 1.2 Outboard Horizontal Stabilizer Parent Vehicle

The OHS configuration was developed to some degree in the 1940s by Chance

Vought's XF5U Flying Pancake, and more recently by Scaled Composites for NASA and

Section 1.2: Background of the Outboard Horizontal Stabilizer Configuration 19

at the University of Calgary where Prof. John Kentfield and Dr. Jason Mukherjee designed

and built a few models. The author has created structural dynamic models of the OHS air-

craft and performed analysis to determine the effects of the structure's flexibility on the

control of the aircraft.

1.3 Thesis Overview

This thesis consists of two basic parts. The objective of the first part is to discuss the com-

ponents and the workings of the PCUAV system. Emphasis will be placed on the author's

contributions to the project, but much will be said about work done by other members of

the PCUAV team. The second part of the thesis will discuss the author's work on the struc-

tural and aeroelastic modeling of the Parent aircraft and the findings from that work.

Chapter two discusses the components of the PCUAV system. This includes discus-

sion of the aircraft developed, the control and navigation system, the communications and

surveillance system, and the reintegration system.

Chapter three focuses on the work done on the project by the author. More detail is

given to the building and testing of the various vehicles.

Chapter four describes the work done on the structural modeling of the Parent,

describing the analysis process and the some of the results from that analysis. Attention is

mostly given to the natural frequencies and mode shapes of the aircraft.

Chapter five discuses the aeroelasticity, flight dynamics, and servo aeroelasticity of the

flexible the OHS Parent aircraft; providing information that could be useful in modifying

the plane's flight controller.

Chapter six provides a summary of the PCUAV system and the structural modeling

done by the author.

20 Chaoter 1: Introduction

Chapter

2PCUAV System

2.1 Chapter Overview

This chapter explains the concepts of the PCUAV system. A basic overview of the sys-

tem's procedures is discussed, followed by a description of the subcomponents of the sys-

tem.

2.2 The PCUAV System Concepts

This section describes the potential roles of the PCUAV system, the process that the sys-

tem's components follow to achieve a successful flight, and some of the key enablers for a

successful mission.

2.2.1 Typical Flight

PCUAV could be utilized whenever there is a need to perform up-close surveillance in

hazardous situations, without risking human life. This could be in a wide variety of con-

texts including collecting soil, air, and water samples from a nuclear waste site; taking

images of a battle zone simultaneously from several altitudes and positions; and locating

targets that may be undetectable from satellites or high altitude UAVs.

21

22 Chanter 2: PCUAV System

In a typical flight, a Parent would be loaded with two Minis, up to six Payload Deliv-

ery Vehicles (PDVs) and MAVs, and enough fuel for a trip of one to two hundred miles,

with a five hour loitering time over the target. Of course range and payload could be

traded for endurance. The Parent would take off from either an airport or an unimproved

landing strip and fly either autonomously or remotely to a predetermined point of interest.

Meanwhile, a communications link would be maintained with a ground operator at the

point of departure. After arriving at the target zone, the Parent would deploy a Mini which

could descend to a lower altitude for surveillance until it ran low on fuel, at which point it

would return for reintegration with the Parent and the second Mini could be deployed. The

first Mini could then be refueled and the process repeated. During the time of the Minis'

operation, the Parent would drop PDVs and MAVs to gather soil samples or provide even

closer surveillance than can be achieved with the Minis. Some of the soil samples might

be collected by a balloon rendezvous with the Mini and brought back to the Parent for

reintegration and transport home. During the whole mission, a communications network

between each of the aircraft would also provide a link back to the ground operator to pro-

cess the data gathered by each vehicle and sensor. When the Parent returned to base, it

would be fueled and readied for a second flight. With the use of just two systems working

concurrently it is feasible to expect a indefinitely sustained presence over a target.

22

Section 2.2: The PCUAV System Concepts 23

PARENT

Figure 2.1 Communications Hierarchy for PCUAV

2.2.2 Key Enablers

Some of the keys to the above hypothetical mission are:

- Building a Parent capable of carrying a large load a long distance,

- Maintaining stability of the Parent when combined with one or two Minis as well as

without them,

- Creating a robust procedure for the system to follow during reintegration to mini-

mize the chance of failure,

- Creating a communications network that is dynamic and dependable between all of

the aircraft.

During the first year of the project, much work was done to size the potential vehicles.

It was hypothesized that a capable Parent aircraft would have a wingspan of about twelve

to fifteen feet and each Mini would have a wingspan of roughly 3 to 5 feet. These projec-

tions were determined assuming that the best possible custom made micro-electronics and

power plants would be utilized. The downside of this assumption is that these things are

expensive and difficult to obtain. Using off the shelf electronics and standard remote con-



trol aircraft engines meant that both aircraft grew considerably and limited the possibility

of completing the desired mission. For this reason, a distinction was made between objec-

tive and demonstration vehicles. The objective vehicles would make use of the best avail-

able products and would be of similar size to the group's original designs. However, the

demonstration vehicles are designed to show that implementing the other keys to the prob-

lem is possible, leaving the problem of creating a high load carrying, long endurance Par-

ent and smaller Minis to future development.

Looking at the key enablers of the mission, it becomes apparent that one of the most

difficult parts to achieve, which has not been investigated in detail before, is the reintegra-

tion procedure. The concept of joining two aircraft in midair is not new. It dates back to

the days of dirigibles that carried biplanes and moves through to the 1960's when work

was done investigating the docking of parasite aircraft such as the Goblin to large bombers

like the B-52. Now, the joining of aircraft is done every day through midair refueling. The

difference in the case of PCUAV comes in the fact that these aircraft are autonomous, with

minimal onboard avionics. The advantages of reintegration extend past the use of the

PCUAV system as a reconnaissance tool. Some examples might include aerial refueling of

UAVs and recovery of soldiers floated aloft by balloons.

2.2.3 Reintegration

To assure that the aircraft will make a rendezvous with a minimum chance of aborting the

procedure, it is important to establish a routine for the planes to follow every time. For this

reason, the rendezvous process was divided into 3 phases. Phase one is designed to bring

the aircraft from arbitrary positions and velocities in the sky to a point where the Parent is

flying about 30 feet in front of the Mini. Phase two brings the Mini into contact with the

Parent, and phase three locks the interface between the aircraft and positions the Mini in a

24

place suitable for transport, such that the combined Mini and Parent have minimum aero-

dynamic drag. During the PCUAV project, the goal is to demonstrate the first two phases,

leaving the third phase for design of the objective vehicles.

In phase one, the Parent is assumed to be circling at a higher altitude, while the Mini is

performing surveillance below. The planes communicate their GPS obtained positions to

each other and the Mini computes a trajectory to follow for rendezvous. This trajectory is

broken into four steps: a climb, a straight leg, a turn toward the Parent's circle, and another

straight leg. The climb brings the Mini to the same altitude as the Parent. The length of the

first straight leg is then computed so that when the Mini performs the turn and second

straight leg at a constant velocity it will enter the Parent's circle 10-30 meters behind the

Parent. In Figure 2.2, the climb, first straight leg, turn, and final straight leg are denoted by

numbers 1, 2, 3, and 4 respectively. MO and PO are the Mini's and Parent's initial posi-

tions. The scale for Figure 2.2 is in meters.

3)600

404)

2 20000

800

600

400

200

0

y axis -200 -200 0 200 400 600 800 1000

x axis

Figure 2.2 Phase One of Reintegration


All of phase one is done under GPS navigation using a proportional navigation algo-

rithm. Onboard GPS data is gathered at a rate of 5 Hertz, and at that same 5 Hertz rate,

points are chosen on the trajectory 100 meters ahead of the current position of the Mini.

The plane flies toward that point until a new point is chosen .2 seconds later. During the

climb and the first straight leg, the trajectory is recomputed at 5 Hz as well. Later, during

the turn and second straight leg, the trajectory is fixed and the velocity is updated instead

so as to align both the planes' position and velocity at the point of rendezvous.

Once the Mini comes to its proper position behind the Parent, phase two begins. Dur-

ing phase two the aircraft navigation is performed using a vision based system. A camera

is mounted under each wing of the Mini at about half span. A cooperative light mounted

on the parent is imaged by the cameras. Using stereoscopic vision the relative position of

the planes can be computed to an accuracy of a few millimeters. The original system

developed in the second and third year of the project relied on two color cameras whose

images were downloaded by a frame grabber to the computer. The images were analyzed

to find the target, which was a large red light the size of a car's headlight. Once the cam-

eras detected the red light the computer used the pixels in the image to determine where

the Parent was.

Infrared Detector

.................

Infrared LEDs Infrared Detector

Figure 2.3 Phase Two Detection and Navigation System

26 Chaoter 2: PCUAV System

Section 2.3: Components of PCUAV 27

This approach and equipment had several problems. First, many things on the ground

are red and create false targets. Second, the range of the cameras for finding the target was

around 10 meters, thus the close proximity required tested the accuracy of the GPS,

increasing the likelihood of a midair collision. Third, the position of the sun had a large

impact on the quality of the images from the cameras. Sunlight, reflected from various

parts of the Parent tended to produce false images and confuse the guidance system. The

final problem was that the update of the frame grabber used was around two seconds when

the whole image was processed. To get around this, the image was zoomed in once the

cameras had a lock on the target. This required the frame grabber to capture many fewer

pixels and the processing time was brought down to 0.2 seconds. Because of these four

problems, an alternative was sought. The cameras were eventually replaced by infrared

detectors that measure the location of the centroid of the light from the target. These elim-

inate the need for the frame grabber. Instead, all that is needed for position processing is

four analog to digital converters to compare the signals from each detector on both axes.

By pulsing the infrared light of the target at 4 kHz the signal is likely to be different from

anything occurring in nature, thus reducing the probability of locking onto a false target

and the possibility of the cameras being saturated by the sun's light. Finally, the range of

the LED system is four times that of the headlight system using the same power output,

even though the surface area of the target light is reduced to 25% of the original headlight.

This means the GPS navigation does not need to bring the planes as close together, and the

drag penalty for the Parent is reduced by the smaller light.

2.3 Components of PCUAV

This section contains descriptions of the components of PCUAV. These include the Parent

and Mini vehicles, Avionics Testbed Aircraft, Payload Delivery Vehicles, the Mini-Parent


Integration Mechanism, the Mid-Air Recovery System, the communications and surveil-

lance systems, and the flight avionics used in the system.

2.3.1 Parent Vehicle

During the first two years of the PCUAV project much was done to choose a design for the

Parent vehicle. A set of requirements was established for the ideal, or objective, vehicle.

These are as follows:

- 100 mile range

- 5.0 hour loiter time

- Autonomous navigation

- Carry a load of two Minis and six Micro vehicles or payload delivery vehicles.

- Capable of performing reintegration with Minis

- Operate from unimproved fields

- Be transportable in a convenient package

Many different concepts were evaluated, most of which used fairly conventional con-

figurations. Some examples are discussed by Sanghyuk Park and Francois Urbain in their

theses, [8] and [10]. As the work progressed, the scale of the Mini continued to grow due

to the need to use large off the shelf electronics. This made the scale of the Parent grow to

the size of an ultralight or small general aviation aircraft, if it was to carry two Minis. As a

consequence, the requirements had to be changed for the demonstration vehicle. It was

decided that the key to the PCUAV concept was to demonstrate reintegration of autono-

mous aircraft. This meant that the requirements for the demonstration vehicle were

changed to the following:

* Be transportable in the MIT Aero/Astro department minivan, see Figure 2.4

28

______________________________________________________________ _____________~~~1 ~ - ~-- - - -


- Fly autonomously

- Loiter for .5-1 hour.

- Maintain speeds suitable for reintegration

- Provide a relatively stationary, stable target for the Mini

- Continue flying after docking with one Mini

Figure 2.4 Parent Aircraft Inside a Dodge Caravan

The team decided that the best option, with the most advantages for reintegration, was

the OHS configuration. Many large scale UAVs are capable of the first five requirements

for the demonstration vehicle, but with a conventional configuration there is concern

about the aerodynamic interference associated with having the tails of both aircraft in

close proximity to each other. With the OHS configuration, the tails of the Parent are well

outboard from the Mini and the interference is minimized. Other advantages of the OHS

are presented in more detail in Section 3.2.1.

2.3.2 Mini Vehicle

As with the Parent design process, in the design of the Mini vehicle there were discrepan-

cies between the objective vehicle and the demonstration vehicle requirements, yet these

differences do little in changing the configuration of the aircraft. However, due to the size

of available electronics, the demonstration vehicle is much larger than the objective vehi-

cle would be. The requirements for the demonstration Mini vehicle are as follows:

- Fly and navigate autonomously

- Maintain a speed appropriate for reintegration

- Maneuver to hit a target on the Parent for reintegration

- Minimize detrimental effects on the flight of the Parent

For these reasons a design was created featuring a pusher propeller, a vertical direct

side force fin, and flaperons that can be positioned both up and down. The pusher configu-

ration decreases the chance of a prop strike when the Mini approaches the Parent from

behind, and also clears an area for a probe to extrude from the nose for reintegration. Hav-

ing a vertical fin over the center of gravity of the aircraft creates the possibility of moving

side to side without yawing or banking when the fin is combined with both aileron and

rudder deflections. Likewise, it is possible to make the plane move up and down with the

combination of flaperons, elevator, and throttle without pitching or changing airspeed.

Figure 2.5 (Left) NGMI, (Right) NGMII

The first Mini was built in the second year of the project. It was made of fiberglass

reinforced foam and had a wingspan of roughly four feet. It was used to do wind tunnel

Chapter 2: PCUAV System30

testing to determine flight characteristics and aerodynamic derivatives to be used in the

control system. This aircraft was later damaged in an office fire, which led to the building

of a second Mini called the New Generation Mini, or NGMI, that is 25% larger than the

first. Later it was found that even this increase in size was insufficient to accommodate the

flight computers, and a third Mini, NGMII, was built 15% larger than the second. These

two vehicles are shown in Figure 2.5. More details of the Mini's construction are pre-

sented in Section 3.3.

2.3.3 Avionics Testbed Aircraft

Testing new avionics and control systems is a risky venture, and can easily lead to crashes.

Realizing that the building time of several months for one Mini or Parent aircraft was

more than the group could afford in the event of accidents, other solutions to testing avion-

ics were investigated. A surrogate aircraft, called the Avionics Testbed Aircraft (ATA),

was found in the Hobbico Superstar 60. This kit plane is available "almost ready to fly"

and only requires a few weeks of work to modify for carrying PCUAV flight computers.

The ATA aircraft is shown in Figure 2.6. Two of these aircraft were built and they both

provided large amounts of insight into what did and did not work. Unfortunately, they both

crashed and were damaged beyond repair. However this was much better than what might

have resulted if these accidents had happened with any of the other aircraft. More about


the work done on the ATAs is presented in Section 3.4.

Figure 2.6 The First Avionics Testbed Aircraft

2.3.4 Payload Delivery Vehicle

During the second year of PCUAV, a payload delivery vehicle (PDV) was developed. This

vehicle was designed to be dropped from the Parent and to deploy ground sensors or

robots. The requirements of the PDVs are as follows:

- Carry a payload the size of a six inch cube and weighing one to two pounds

- Impact the ground at less than 16 feet per second

- Withstand a 30g impact load

- Land within 45 feet of a target

- Be no larger than 6 inches by 6 inches by 20 inches to fit inside the Parent

- Weigh less than 3.5 pounds.

- Cost less than $5000 each for expendability

" Drop from approximately 6000 feet.

Several configurations were considered including a parafoil, a winged body, a cylin-

drical body, a lifting body, and a rotating wing controlled descent aircraft. Out of these

options, a lifting body concept was chosen because it offered the most control and the least

32 Chapter 2: PCUAV System

33

susceptibility to wind due to its relatively high descent speed. A concept vehicle (see Fig-

ure 2.7) was built using a foam core that was covered in fiberglass, carbon, and kevlar. The

body is a NACA airfoil with three symmetric tail surfaces for stability and control. This

vehicle met or exceeded all of the requirements. It was tested by dropping it from the top

of a 60 foot building onto asphalt. The aircraft demonstrated stable flight and descended at

its design angle of decent of 60 degrees.

Figure 2.7 Payload Delivery Vehicle

2.3.5 Mini-Parent Integration Mechanism (MPIM)

The Mini-Parent Integration Mechanism, or MPIM, is the physical mechanism that forms

the connection between the Mini and Parent. Many ideas were explored for the MPIM.

The original idea for the objective vehicles was an extending arm that protruded at an

angle up and behind the Parent. (See Figure 2.8) The arm had a grabbing claw that would

clamp down on a catching ring on the nose of the Mini and then the arm would retract and

bring the Mini to the top of the Parent where it could be locked down for transport. This

system is somewhat complex and heavy and requires high navigational accuracy because

SEE

Section 2.3: Components of PCUAV


of the limited target size.

Figure 2.8 Original MPIM Design

Another idea was a drogue trailing from the Parent and a probe on the Mini similar to refu-

eling aircraft. The line would be reeled in after contact was made between the planes. The

problem seen with this system is the possibility of a pendulum effect between the two air-

craft when the line between them becomes short.

Figure 2.9 Parent Aircraft with MPIM Attached

Because the goal of the PCUAV project is to demonstrate phases one and two, it is

only necessary to make contact between the two planes, meaning they do not need to be

physically attached to each other in flight. A system was designed and constructed to pro-

34

Section 2.3: Components of PCUAV

vide a target without a locking mechanism. (See Figure 2.9) This system is a three foot

high truss that is attached to the top of the Parent's fuselage. The truss is made of 5/8 inch

carbon/epoxy tubes that are wrapped by 1/16 inch balsa sheet that provide an aerodynamic

shape to reduce drag. The target light sits on the top of the truss, and a drogue "catcher" is

attached six inches below the light. The drogue is a four inch diameter, eight inch long

fiberglass cylinder that expands into a conical net with an eighteen inch diameter. A wood

block was inserted into the mouth of the cylinder. The block has an oblong hole in it that

aligns with the probe attached to the front of the Mini. This hole restricts the rolling of the

Mini, keeping the wings parallel to the wings of the Parent. The drogue is attached to the

carbon truss with a steel axle that allows the Mini to pitch against the resistance of a

spring. Because the Mini can apply the most moment on the truss through pitch, a spring

attachment lessens the load on the truss. At the same time, both roll and yaw are con-

strained.

A rudimentary spring loaded locking mechanism was built inside the cylinder as a

concept for phase three. (See Figure 2.10) The probe on the Mini has a knob that slides

between two arms that are snapped behind the knob by springs. When it is time to disen-

gage the two planes, a servo spreads the two arms, releasing the Mini to fall back.

Locking Mechonism

Figure 2.10 Detail of MPIM

35

2.3.6 Mid-Air Recovery System

During the course of a typical mission it may be desirable to recover something from the

ground, such as soil samples or valuable equipment, and, on a large scale, possibly even

people. A Mid-Air Recovery System, or MARS, was developed in the PCUAV project for

such a purpose. The system consists of a balloon to carry the desired package, an RF trans-

mitter attached to the balloon, and a directional receiver onboard the rendezvousing plane.

To collect a soil sample, a PDV would deliver a collector from the Parent to the desired

position on the ground. A sample would be taken by the collector, the balloon would

inflate, and the collector would rise to the altitude of the Mini, where an RF transmitter

would send an omnidirectional signal. The receiver on the Mini would find the bearing

toward the transmitter and steer the Mini towards it. The Mini would then fly through the

cable connecting the sample to the balloon and retrieve the sample at the wingtip while

cutting away the balloon.

A transmitter and a directional receiver were designed and constructed. The receiver

uses four antennas oriented in a pyramidal orientation as seen in Figure 2.11. The strength

of the signal is compared between each antenna. When the strength is equal on each, the

transmitter is directly ahead of the Mini. The system could be calibrated to determine the

angle toward the transmitter based on the signal from each of the antennae. This system

36


could also prove useful for reintegration with the Parent as a supplement to both the GPS

and vision systems.

Figure 2.11 MARS Directional Finder

2.3.7 Communications and Surveillance

Work was done by Alexander Olmenchenko on the communications and surveillance

aspect of PCUAV [7]. The surveillance system consists of a computer stack separate from

the one used for navigation, a video camera, and a WLAN system. The system was flown

in both the ATA and NGMI. The first concept demonstrated was the ability to take video

images from the air and transmit them via WLAN to a laptop computer on the ground. The

pictures are recorded, compressed into JPEG form, and transmitted to the laptop at a rate

of one frame every two seconds. The second concept used a camera placed on the ground,

which transmitted images to a UAV overhead which relayed them to the laptop. At the

same time, the laptop operator was able to move the ground camera through commands

sent back through the aircraft. The third concept demonstrated was the ability to send and

receive images from both the airplane and the ground cameras at the same time. The next

concept to be demonstrated will be putting the ground camera on a rover to show that it

can be operated by the laptop with a signal relayed through the airplane. The final surveil-

lance concept to be demonstrated will combine all of the previous concepts with GPS so

that the airplane can be made to orbit a moving rover based on its GPS position. This

would allow the user to view both an image from the rover and an image from the aircraft

of the rover at the same time. Figure 2.12 shows the rover and ground camera used for sur-

veillance.

Figure 2.12 Rover with Surveillance Equipment on Top

2.3.8 Flight Avionics

The three PCUAV flight vehicles utilize similar avionics packages. They all have com-

puter stacks, RC receivers, RF transceivers, GPS, air data sensors, relay switches, gyro-

scopes, and accelerometers. In addition, the Mini has two infrared detectors and their

related electronics for use during Phase two of reintegration. Figure 2.13 shows how the

38 Chapter 2: PCUAV System


avionics components are configured in the Mini.

ElectricalRadio

.... RS232

Figure 2.13 NGMII Flight Control Avionics

The CPU used in the computer stack of each plane is a 233 MHz processor made by

Real Time Devices USA, Inc. that runs DOS. This computer interfaces with all of the ana-

log flight sensors through a Real Time Devices 16-bit databoard. The system is powered

through a power board made by TriM that provides +/-5V and +/-12V. Programs are

downloaded, and flight data uploaded through a CM312 utility board from Real Time

Devices. This module provides communication between the computer and the GPS, RF

transceiver, and SBC2000s.

The SBC2000s, made by Micro Pilot, provide an interface between the aircraft's ser-

vos and either the pilot or the flight computer. In pilot-in-command mode they read the

pilot's inputs through the RC receiver #1 and send the appropriate pulse width signal to

each of the servos. In computer-in-control mode, the SBC2000s produce the same signals

based on input from the computer. The computer creates this input based on several


40 Chanter 2: PCIJAV qvot-m

sources of information. The primary source of information during Phase one of reintegra-

tion comes from the GPS. The system used for PCUAV is an All-Star GPS from BAE Sys-

tems, Canada. It provides information at a rate of 5 Hz. Because PCUAV's aircraft are

relatively small and have small time constants of motion, they require higher accuracy of

position from the GPS than is available in order to achieve navigation based solely on

position. For this reason, velocity and acceleration are measured and combined with fil-

tered position data to aid in producing smooth, stable guidance, navigation, and control of

the vehicles.

Other key information for navigation comes from a set of rate gyros, an accelerometer,

and pressure sensors. The rate gyros are made by Tokin. These gyros are capable of detect-

ing deflection rates of up to 300 degrees per second with a resolution of 1 degree per sec-

ond. They do have a large drift rate of about 1/3 of a degree per second, and are

susceptible to vibration noise. For the purpose of redundancy, there are six gyros in total,

two per axis. The accelerometer, made by Crossbow, provides information in three direc-

tions with a range of 4g's and an accuracy of 0.005g's. Pressure sensors from Omega pro-

vide static and dynamic pressure for both altitude and airspeed information. The airspeed

sensors for both the NGMII and the Parent were calibrated in the MIT Wright Brothers

Wind Tunnel.

During Phase two of reintegration, the primary information role of GPS in the Mini is

superseded by the vision system described in Section 2.2.3. This system consists of two

infrared detectors, made by Pacific Silicon Sensors, Inc. The signal from each camera is

sent through an analog to digital conversion board and combined to determine the position

of the aircraft relative to the target.

40

The information from all of the sources described above is communicated between the

two vehicles as well as with the ground station through RF transceivers made by Max-

Stream Inc. These have a range of about 1.5 miles with the antenna chosen for PCUAV.

The aircraft are controlled by the pilot through standard RC radio gear made by Fut-

aba. The copilot uses a similar set of radio gear to turn the computer on and off and switch

between pilot in control and computer-in-control modes. The range of RC equipment can

be affected by the computer's electromagnetic interference, reducing the range of the

pilot's control in a field test by as much as half when the computer is switched on. Manu-

facturers other than Futaba were also tested, but Futaba was found to produce the best

range. Efforts were made to shield the computer's noise, and the best range was achieved

when the antennas from the pilot's and copilot's receivers were separated as far as possi-

ble. The other choice in radio gear is between FM and PCM. This choice has little effect

on range, the main difference being how the aircraft behaves when signal is lost. In FM,

the servos start to jitter when the signal weakens, whereas they hold their last known posi-

tion in PCM mode. Even though the FM jittering makes signal loss easier to detect, PCM

was chosen for PCUAV because other onboard electronic devices are adversely affected

by jitter noise.

The final, and one of the most important components of the onboard avionics, is a six

channel relay switch. During flight, the pilot may choose to fly in one of two modes, nor-

mal mode and safe mode. In normal mode the pilot's signal from the receiver is routed

through the relays to the SBC2000s and then to the aircraft's servos, which allows the

computer to control the plane. When the pilot switches to safe mode, the relays bypass the

SBC2000s, providing a hardwire connection from the receiver to the servos. This mode

could save the plane in the event of a computer malfunction. Unfortunately, this mode

switch was a factor in the destruction of the second ATA. The plane had stalled while


under computer control and went into a steep dive. The pilot switched to safe mode, but

applied full up elevator before making the switch, the result was that the plane experi-

enced an instantaneous 7g pull up that snapped the wing at the root. Afterwards, a proce-

dure was created for switching to safe mode. Before making the switch, the pilot centers

all of the control surfaces so that when the switch happens minimum stress is induced

before any maneuvering is attempted.

Three sets of avionics were built for PCUAV. One for the ATA and NGMII, one for the

Parent, and a third that was used in the lab to test flight codes. Because the avionics can be

difficult to remove from the aircraft, and there is a chance of disturbing connections while

moving the avionics, it was important to have a separate and identical set in the lab for

testing. This set was connected to a hardware in the loop simulator which consists of ser-

vos similar to those on the aircraft that are linked to potentiometers that provide feedback

to the simulator computer. In this way all flight codes could be tested on the ground before

ever being put in the aircraft.

The ATA and NGMII have virtually identical avionics while the Parent's is similar, but

splits the controls into three parts. The Parent was built with one receiver in the fuselage to

control the throttle and nose gear and one receiver in each wingtip to control the rudder,

elevator, and aileron on each side. This shortens the length of the servo wires for these

controls to reduce the RF interference associated with long wires. To achieve autonomous

flight, a relay, a SBC2000, and a battery must be added to each of the three receivers. A

serial link along the underside of the wing connects the wingtip avionics with the central

computer in the fuselage, which differs from the Mini's avionics only by the replacement

42

of the rate gyros with an inertial measurement system built by Crossbow. See Figure 2.14

for more detail on the Parent's Avionics.

GS treceiver

RF transceiver

electrical

radio lapto_-,- radi RF transceiver 6 - lpo- serial pilot co-pilot

Figure 2.14 Parent's Avionics Structure

2.4 Chapter Summary

Chapter two presented a background of the PCUAV System. Section 2.2 presented a

description of the mission of PCUAV as well as some of the key enablers of the mission.

Special attention was given to the parts of the mission in which demonstration is desirable,

particularly reintegration, which was separated into three phases. Phase one being oper-

ated under GPS navigation, Phase two under guidance of a vision system, and Phase three

making a solid physical connection between the two aircraft. Section 2.3 went on to

describe the components of the PCUAV system, stressing those components the author

was most involved with in the project. These include much of the design and building of

the Parent and Mini vehicles, the Avionics Testbed Aircraft, the Payload Delivery Vehicle,

Section 2.4: Chapter Summarv 43

Chapter 2: PCUAV System

the Mini-Parent Integration Mechanism, and the various avionics utilized in each of the

flight vehicles.

44

Chapter

3UAV Building and Testing


This chapter discusses in depth the vehicle designs, including their dimensions, structural

materials selection, and techniques employed by the author and other team members in

building these aircraft. A brief commentary on flight testing for each vehicle is also pre-

sented.

3.2 OHS Parent Vehicle

This section describes the Outboard Horizontal Stabilizer configured Parent vehicle built

by the author and Francois Urbain. The advantages and disadvantages of the OHS config-

uration are discussed, as well as the dimensions and performance of the aircraft. The con-

struction of the three main components of the aircraft, namely the wing, the tails, and the

fuselage is also described.

3.2.1 Advantages and Disadvantages of the OHS

The key demonstration goal for PCUAV is reintegration of the Mini and OHS vehicles.

For this reason, both vehicle designs were created with that purpose in mind. The unique

tail configuration of OHS aircraft provides a maximum amount of clear space behind the

45

Chapter 3: UAV Building and Testing

fuselage. This keeps the tail surfaces away from any flow disturbances from the Mini and

reduces the chances of a collision between the Mini and the Parent's tail when compared

to a conventional configuration. Analysis was done using computational fluid dynamics to

determine that the OHS Parent would remain controllable even after a losing one tail sec-

tion as a result of a midair collision. The other advantage of the OHS configuration is the

fact that the horizontal stabilizers are flying through the upwash of the wingtip trailing

vortex. (see Figure 3.1) The lift is thus distributed to the tail as well as the wing. This

means more lift is possible with the same wing area as a conventional configuration where

the tail is in the downwash of the wing's vortex and usually pushes down, requiring more

lift from the wing.

-2j

0

02

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2

y/b

Figure 3.1 Vortex Induced Angle of Attack at Tail Position

The main disadvantage of the OHS aircraft comes from the moments generated by the

lifting tail on the wingtips. The wing must be stiffer and stronger in torsion than wings on

conventional aircraft. This can be achieved most easily with a thick airfoil of relatively

low aspect ratio.

46

3.2.2 OHS Parent Design Process

After deciding on the configuration of the Parent, the next step was to design and size the

specific aircraft applicable for the project. A few things were drivers for the final design.

First, the plane needed to be large enough that the Mini would have little trouble clearing

the Parent's tails with its wingtips. For this reason the wingspan of the Parent was chosen

to be roughly twice that of the Mini. Second, the wing had to be stiff enough to resist the

torsional moment from the tail booms. This led to an aspect ratio of 8.0 with a NACA

2412 airfoil. At the same time, the wing needed to fit in the back of a mini van. For this

reason the wing was built in two pieces, each about 7.5 feet long. The tail booms are also

detachable to enable transport, as shown in Figure 2.4.

The tail volumes were chosen [9] to provide stability similar to general aviation single

engine aircraft. These are calculated by (eq. 3.1) and (eq. 3.2), where vvt is the vertical

tail volume coefficient and Vht is the horizontal tail volume coefficient. The other sym-

bols are listed in the list of symbols and acronyms on page 6.

V = Lvt XSvt (eq. 3.1)vt bw x S

Vht Lht xSht (eq. 3.2)c, X SW

The values picked were vvt = 0.04 and vht = 0.7. The tail length was set by the avail-

ability of carbon tubes. It was desirable to use a tube that was of a diameter similar to the

thickness of the wing so it could blend smoothly into the wingtip. The tubes chosen have

diameters of 1.125 inches and are sold in six foot lengths, thus fixing the length of the tail

and the tail area for the chosen tail volumes. A NACA 0012 airfoil was picked for the tail

airfoil since it is a common, well proven airfoil shape that is relatively easy to build. A

Section 3.2: OHS Parent Vehicle 47

more efficient design, such as those done at University of Calgary and described in [3] and

[5], would use a non symmetric lifting airfoil for the horizontal tail.

The size of the fuselage was largely driven by the size of the avionics package which

requires a 1 Ox 1 Ox 18 inch box. Attached to the front of this box is a section large enough

for the fuel tank, tapering to the engine. To the rear of the box is a tapered section to

reduce the possibility of laminar airflow separation. The initial engine chosen was an O.S.

max- 160FX which is the largest glow powered, two stroke, single cylinder engine built by

O.S. It produces 3.7 horsepower. As described in Section 3.2.4, this engine was later

replaced with a more powerful Moki 2.10 engine.

The final dimensions of the Parent are presented in Table 3.1.

Table 3.1 OHS Final Dimensions

Wing Span 180 in. Length 103 in.

Wing Chord 21.25 in. Height (w/o truss) 45 in.

Wing Area 3825 in.2 Height (w/ truss) 56 in.

Tail Span 233 in. Fuselage Dim. 10.5 x 10.5 x 43 in.

Horiz. Tail Area 738 in.2 Empty Weight 41.1 pounds

Vert. Tail Area 426 in.2 Airfoil NACA 2412

Avg. Tail Chord 12 in. Engine (hp) Moki 2.1 in. 3 (4.9)

3.2.3 OHS Construction

Construction of the OHS Parent began with the wings. Each of the two main wing spar

halves were made from two pieces of 1/4 inch plywood sandwiching three pieces of 1/4

inch balsa. A layer of 2-ounce fiberglass was epoxied between each layer of wood at a 45

degree orientation to help strengthen the spar in the shear direction. The two spar halves

overlap at the center by ten inches, with the plywood pieces interlacing each other. Two 1/

48 Chapter 3: UAV Building, and Testing,

4-20 steel bolts hold this center section together with a two degree dihedral angle. Using a

milling machine, the spars were beveled on the top and bottom to match the contour of the

airfoil. Finally, seven layers of unidirectional carbon fiber were laid on the top and bottom

faces of the spars to support the bending load of the wing. (see Figure 3.2) Despite the fact

that the length of the spar is only 40% of wingspan, and it is only 1.5 inches wide, it repre-

sents 30% of the total weight of the wing and provides most of its bending strength. A rear

spar of half inch balsa runs the full length of the trailing edge. It is reinforced with 1/8"

plywood at the wing root, where the two halves are bolted together.

UnidirectionalCarbon Fiber

Balsa

%" Plywood

Fiberglass BetweenWood Layers

Figure 3.2 Parent's Spar Detail

While the spar was being manufactured, a foam core was cut for the rest of the wing

using a foam cutter. Because of the size of the cutter, each wing had to be made from three

sections of foam epoxied together. The inboard section was notched to accommodate the

spar. The leading and trailing edges were cut off and later a solid balsa leading edge and a

hollow balsa trailing edge were glued on. The skin of the wing was made from 1/16 inch

balsa sheets with the grain running along the length of the wing to resist bending. The

49

50

balsa was reinforced with a composite layup on the internal side. A layer of 4-ounce fiber-

glass was laid along the full span at a 45 degree orientation to provide torsional stiffness,

as well as two extra layers at the wingtips to give extra stiffness where the tail booms con-

nect to the wing. One layer of unidirectional carbon and one layer of carbon fiber cloth

oriented parallel to the span was laid over the area of the spar. These carbon layers were

cut into a diamond shape to smooth the transition of the bending stress from the wingtips

into the spar and then again into the center section of the spar. (see Figure 3.3) Once the

fiberglass and carbon was laid on the balsa sheeting, the skins were placed on the top and

bottom of the foam cores with the spar halves epoxied in them. The entire wing was then

placed in a vacuum bag and put under pressure. Because the composite layup was put on

the internal side of the balsa, little sanding was required to produce a smooth finish after

the wing came out of the vacuum bag.

Heavy Carbon Fiber @ 45 deg Unidirectional CarbonFiber

Spar

4-oz Fiberglass @ 45 deg Extra FiberglassLayers

Figure 3.3 Parent Wing Composite Layup

When building the tail section it was important to keep the weight as low as possible to

reduce the moment on the wingtip produced from the inertia of the tail during landing. For

this reason, the tail was built with balsa in lieu of a foam core fiberglass surface. Eighth

inch balsa ribs were glued to a 1/4 inch balsa spar and the leading 25% of the tail was

sheeted with 1/16 inch balsa. The horizontal and vertical surfaces were fiberglassed to a


Section 3.2: OHS Parent Vehicle

solid piece of balsa that had been sanded to a streamlined shape and notched to fit to the

end of the carbon tail booms. The horizontal surfaces were given eight degrees of dihedral

to avoid ground contact in a wing low landing. (see Figure 3.4) The tail booms were glued

into a wingtip pod that was bolted into two oak blocks epoxied into the front and back of

the wing's tips.

Figure 3.4 (Left) Author with Parent's Tail, (Right) Parent's Fuselage Frame

The last component of the aircraft to be built was the fuselage. This was built using 5/

8 inch carbon tubes to form a truss structure. (see Figure 3.4) The carbon tubes were held

together by epoxy embedded with shreds of fiberglass. This truss connects all of the high

stress locations; which are the engine and nose gear mount, the front main spar mount, the

rear spar mount, the main landing gear, and finally the attachment points for the MPIM.

This structure provides excellent strength with little weight. The frame was wrapped with

1/4 inch balsa to give it the appropriate shape. The bolts that hold the main spar together

also bolt into the frame, while the rear spar is rubber banded to the frame. The main land-

ing gear were bolted into an oak block, and the firewall was made from 1/4 inch plywood.

51

52

3.2.4 OHS Testing and Updating

This section describes the testing that was done on the OHS aircraft as well as changes

that were made due to the results of that testing. Included are structural testing and both

taxi and flight testing.

3.2.4.1 OHS Structural Testing

Structural tests were done on the wing for strength in the bending, torsional, and longitudi-

nal directions. The bending load was tested to the same standard that the Federal Aviation

Administration requires for normal category aircraft, 3.5g's, plus a factor of safety of 1.5.

Linear deflection of the wingtips was observed all the way to 5.5g's at a rate of 1/8 inch

per g-force. The OHS configuration actually has an advantage over normal aircraft struc-

turally when it comes to bending moments on the wing. Because the mass of the plane is

not concentrated at the wing root, but is also distributed to the wingtips, the bending

moment actually switches direction at about half span, leaving a point where there is actu-

ally no bending stress. (see Figure 3.5) This point happens to be close to the end of the

spar, which decreases the adverse affect of having a discontinuity in the structure there.

The overall effect is that the stress at the wing root for the OHS is about half what it would



be in a standard configuration.

Bending Moment3000

Standard Configuration-- OHS

2500 ---------- - ------------ ------------ ------------- --------------

2000 --------- --- ---------------------------- ------------- --------------

C 1500 -------------- ---------- ----------- ----------- -------------

E 1000 ------- - ---------- --------------- -- -------------

500-------------

0 -------- -- ------- -- --- -- -- - ----

-5000 20 40 60 80 100

Span (in.)

Figure 3.5 Bending Moment in Parent's Wing

The torsional stiffness of the wing was tested by applying weight to the tails up to 5g's.

Again, linearity was observed in the deformation.

The last direction that the wing's strength was tested was in the forward longitudinal

direction. This is important for loads seen at high angle of attack, when the lift has a load

component parallel to the wing's chord. The resulting moment can be up to 14% of the

total bending moment on the wing. This load was tested, and no deflection was observ-

able.

3.2.4.2 OHS Taxi and Flight Testing

Once the aircraft was completely built and assembled, taxi tests were done. It was imme-

diately obvious that the landing gear performed less than satisfactorily. The nose gear

oscillated side to side as much as four or five inches even at walking speeds. The main

gear oscillated front and back at higher speeds as well. This problem was fixed by brazing


steel rods to the sides of the landing gear to increase the appropriate moments of inertia.

Tests were performed on both a hard surface and a grass surface. On the hard surface, the

tail booms oscillated with moderate amplitude. Even so, the elevator still demonstrated

some authority and could lift the nose while taxiing at about 50% of takeoff speed. How-

ever, on the rougher grass surface the tails oscillated at high frequency and amplitude, and

the engine struggled to get the plane to high speeds. For this reason operations of the Par-

ent were limited to hard surface runways only.

During the first takeoff of the Parent the tails oscillated at slow speed, but gradually

grew stiffer as speed increased until there were no observable oscillations at takeoff speed.

The aircraft flew smoothly without any undesirable flying characteristics. It was some-

what slower than expected, but had no trouble climbing, and once was glided to the run-

way, after the engine stalled, with a shallow glide angle. (See Figure 3.6 to see an example

of engine out flight.)

Figure 3.6 Second Landing of OHS Parent

For the second flight test, the MPIM was attached to the top of the plane. (see Figure

2.9) This produced significant drag and required full power to keep the plane in the air.

The engine stalled at one point, resulting in a rough landing because the plane's glide ratio

54

had decreased significantly. The MPIM was later better streamlined and the engine was

replaced with a Moki 2.10 that provided 30% more horsepower.

During a third flight test these changes proved that the plane flew reliably and with

performance adequate for reintegration even with the MPIM attached. The flight speed

envelope was measured to range from 18 to 28 m/sec, which is comparable to that of the

Mini. Another problem was observed during this flight test. The pilot found that the plane

would not pitch down from level flight at high speeds even when full down elevator was

applied. At first this was attributed to aeroelastic effects. With OHS aircraft, when the lift

is increased on the tail to pitch the plane down, the lift from the wing decreases, which

decreases the strength of the wingtip vortex, lessening the lift from the tail. This makes for

a very stable system, but also introduces a lag into the elevator effectiveness. This phe-

nomenon should be equally present in both up and down pitch however, which was not

observed in the case of the Parent. The solution was found in the strength of the elevator's

servos. The servos that were being used were the same as what had been used on all of

PCUAV's aircraft, Hitec HS-85MG Mighty Micros. The servos did not provide adequate

torque for surfaces as large as the Parent's elevator. Because the Parent's trim condition

required a few degrees of down elevator, the servos were already being worked to keep the

plane level, thus the asymmetric effectiveness of the elevator. These servos were replaced

with high performance Futaba S9402 servos that delivered 40% more torque.

During the fourth flight test the Parent operated autonomously under computer con-

trol. The plane successfully flew continuous circles, maintaining altitude and position

within half of a wingspan. This was done with winds of 12 miles per hour with even

higher gusts. The control system for the aircraft had been designed for winds only up to 10

miles per hour, so better results are expected with better conditions. In addition, the pres-


56

sure sensor for the altimeter malfunctioned during flight, which meant that all altitude

information came from GPS. The accuracy of this information is only +/- 5 meters.

3.3 Mini Vehicle

This section discusses the design, construction, and testing of the three Mini vehicles and

the continuing evolution of its design. The advantages and disadvantages of the design are

presented. Special attention is given to the construction of the NGMII, which the author

was primarily responsible for building.

3.3.1 Mini Design Process

Like the Parent, the Mini was also designed primarily to perform reintegration. This meant

that the objective vehicle needed to be maneuverable, yet stable, and large enough to carry

the avionics package, yet small enough to be carried by the Parent. A few designs were

proposed during the first year of the project, but they all shared a few common things. All

of the ideas incorporated a pusher propeller to keep the prop away from the Parent, and all

had extra vertical fins to produce sideways movements. Some had the extra fins at the

nose or on the wingtips and others over the center of gravity. Some were low wing aircraft,

with the idea of blending the wing into the wing of the Parent after docking, while others

were high wing planes with the idea of blending the fuselages of the planes together with

the wings arranged in a tandem position to help provide lift to the Parent. All of the initial

designs had a wingspan of less than 60 inches, which may be appropriate for an objective

vehicle. For a demonstration vehicle however, it is only necessary for the Parent and Mini

to make contact, and not necessarily have the Parent carry the Mini, so the size of the Mini

is less restrictive. The first Mini built was more of an objective vehicle in its size and lay-

out. It was a high wing pusher with a vertical surface over the center of gravity. The

Chap~ter 3: UAV Building and Testing

Section 3.3: Mini Vehicle 57

NGMI was designed to be a larger demonstration vehicle. It had the same layout as the

Mini, with the wing and tail being scaled up 25%, and the fuselage made large enough to

theoretically accommodate the flight avionics. However, the team had been optimistic

when sizing the avionics, forcing the building of the 15% larger NGMII.

Because the primary mission of an objective Mini is to loiter over a target, the wing

was designed as a GM-15 airfoil, who's drag bucket fit the mission profile well. A rela-

tively high aspect ratio of 9.1 was chosen for aerodynamic efficiency, as well as Hoerner

wingtips to reduce induced drag. The tail surfaces of the Mini were sized for

V = 0.04 and Vht = 0.54. This makes the Mini a little less stable longitudinally than the

Parent, but still within the bounds set by Raymer in [9].

As was mentioned above, the Mini configuration has advantages from the pusher pro-

peller and the extra vertical fin. There are also some disadvantages associated with the

Mini's design. First, because the engine is behind the wing, it is difficult to place the cen-

ter of gravity far enough forward. It requires either a very long nose or a large nose weight.

Fortunately for PCUAV, the avionics provide nearly enough weight by themselves to bal-

ance the NGMII, so only a small amount of extra ballast weight is required. Second, the

position of the pusher prop limits the amount the plane can rotate during takeoff and land-

ing. This problem could be alleviated somewhat if the main landing gear was both length-

ened and moved rearward. Finally, as will be discussed in Section 3.3.3, the tail booms are

long, giving the tails a lot of mechanical advantage on the attachment point on the fuse-

lage.

3.3.2 Mini Construction

The first Mini was built with a foam core wing, tail, and fuselage; with a few layers of

fiberglass on each. The plane never flew but was used in wind tunnel tests to determine its

flying characteristics. Later the plane was destroyed in an office fire.

The second Mini, also called the New Generation Mini or NGMI, was 75% complete

when the author joined PCUAV. The fuselage was made of 3/32 inch plywood reinforced

with fiberglass. The tail was made of balsa ribs and connected to the fuselage by 3/8 inch

carbon tubes. The wing had a solid plywood spar in a foam core and was sheeted with

balsa and several layers of fiberglass. This resulted in a relatively heavy fuselage and wing

and drove the wing loading to 75% higher than aircraft of similar size. The plane was

powered by an O.S. 0.61FX engine. The NGMI flew well when on long paved runways,

but had a high speed and a limited load capacity. For this reason a second, larger, lighter,

and more powerful NGMII was built.

Figure 3.7 Avionics Inside NGMII Fuselage

The fuselage construction of the NGMII drew from that of the Parent. A truss structure

was made from 3/8" carbon tube connecting the engine, landing gear, and wing. This truss

was sheeted with 3/32" balsa. The nose bulkhead and engine firewall are plywood and the

main landing gear is bolted into an oak block. The nose cone was formed from a shaped

block of foam covered with six layers of four ounce fiberglass. Most of the foam was

removed, leaving a hollow area for a battery compartment surrounded by about an inch of

58 Chapter 3: UAV Building and Testing

foam for impact resistance. The nose was hinged on the left side of the fuselage so it is

able to open for easy access to the batteries much like some military cargo aircraft. The

locking mechanism on the right side can be seen in Figure 3.7.

The tail of the NGMII was made of balsa stick with a bass wood spar and balsa sheet-

ing on the leading 25%. It is attached to the fuselage by 3/8" carbon tubes that are braced

both at the firewall and the landing gear.

1/16' Balsa Skin

End Grain Balsa Core1/8' Balso Rib Carbon Spar Caps

Carbon Web

Figure 3.8 Cross Section of NGMII Wing

Since most of the weight of the NGMI is in the wings, most of the effort to reduce the

weight of the NGMII went into its wing. Instead of using a foam core, this wing was

"built-up." It has a nearly full span spar, balsa ribs, and full wing balsa sheeting. Each spar

cap is made from two half inch wide carbon laminate strips that taper from .06" thick at

the root to .0 14" at the tip. The spar web was made from half inch end cut balsa as the core

material and a 0.5" by 0.03" carbon shear web along the leading edge of the spar. (See Fig-

ure 3.8) These parts were epoxied together and vacuum bagged with two degrees of dihe-

dral and a two inch overlap of the carbon layers at the root. The 1/8" balsa ribs were split

into front and rear pieces and then glued to the spar. Balsa leading and trailing edges were

added and the whole wing was sheeted with 1/16" balsa whose grain is oriented in the

Section 3.3: Mini Vehicle 59

spanwise direction for bending stiffness. See the full set of construction plans in Appendix

A and the final dimensions in Table 3.2.

Table 3.2 Final Dimensions of NGMII

Wing Span 100 in. Length 63 in.

Wing Chord 11 in. Height 21 in.

Wing Area 1070 in.2 Fuselage Dim. 10.5 x 7.25 x 40 in.

Tail Span 30 in. Empty Weight 15 pounds

Horiz. Tail Area 187.5 in.2 Airfoil NACA 2412

Vert. Tail Area 130 in.2 Engine (hp) O.S. 91FX (2.8)

Avg. Tail Chord 6.25 in.

3.3.3 Mini Testing

The first Mini was tested in the Wright Brothers Wind Tunnel during the first year of

PCUAV. Aerodynamic derivatives were found by measuring forces and moments in all six

degrees of freedom while varying control deflection, airspeed, angle of attack, and side-

slip. This data was used to develop the control systems for the subsequent Mini vehicles.

NGMI was first flown in the second year of PCUAV. It flew stably under remote pilot

control and was responsive to controls in all directions. It was heavy however and required

long smooth runways for operation. Later, after the NGMII had been built, NGMI's engine

was replaced with an O.S.max .91FX and the landing gear were upgraded, giving it

enough power and robustness to be useful. It was later used to test the surveillance and

communications systems described in Section 2.3.7, which could fit in the plane's rela-

tively limited cargo space.

Before its first flight, NGMII's wing was load tested to 3.5g's. Sandbags totalling 66

pounds were laid in an elliptical distribution along the bottom surface of the inverted

wing. Linear deflection was observed throughout the loading process, and the wing

60 Chapter 3: UAV Builind tQin

returned to its undeformed state. The wing was also tested in the longitudinal direction,

putting sandbags totalling four pounds on the wingtips. No deflection was observable.

The NGMII first flew in the third year of PCUAV. On its maiden flight, problems were

encountered with the elevator control. The aircraft went into a shallow dive and the pilot

had to use full up elevator and full elevator trim to recover. The joints between the tail

booms and the fuselage were reinforced to reduce the amount that the booms flexed, but

the elevator effectiveness was still not desirable. There was a lag of about half of a second

between the elevator command and the change in pitch due to the flexibility of the booms.

Thin strips of carbon, one inch wide, were fiberglassed edge-on along the length of the tail

booms, increasing their moments of inertia by 215%. This stiffened the booms sufficiently

for responsive flight. The Mini was then used to demonstrate autonomous flights, flying

prescribed circles and simulated paths for phase one of reintegration. These tests were suc-

cessful, with the plane remaining within 4 meters of its prescribed position at all times,

even in the presence of crosswinds and gusts.

3.4 Avionics Testbed Aircraft

This section discusses the avionics testbed aircraft or ATA. Included are the reasons for

the ATA's existence, the work done on the ATA, and the final design of the aircraft.

3.4.1 Advantages and Disadvantages of the ATA

The Avionics Testbed Aircraft was necessary to reduce the risk involved with flying

autonomously. Because the aircraft was relatively cheap and easy to construct, it did not

matter as much if it was destroyed in a crash as if a Mini or Parent vehicle crashed. One

disadvantage of the ATA was that it had a puller propeller, which was undesirable for dem-

onstrating reintegration. The second disadvantage was that the original kit plane was

Section 3.4 Avionics Testbed Aircraft 61

62

designed to be a five pound aircraft, and, when loaded, it weighed three times that amount.

This made the wing loading and structural stress high. Two ATAs were built and flown

during the course of PCUAV. Both were destroyed in crashes. The first was overloaded

and stalled on takeoff. The second plane had extended wings and a larger engine to deal

with the large load, but the wing proved to not be strong enough when it snapped in half

during a 7g recovery from a dive as described in Section 2.3.8.

3.4.2 Work done on ATAs

The ATAs were modified Hobbico Superstar 60's. These kit planes come 90% complete

out of the box and only require a few days work to finish. However, the modifications

done for the ATA took a few weeks. First the fuselage center section was removed and

replaced with one large enough for the avionics package. This fuselage section was built

from bass wood sheeted with balsa. The wings were extended by five inches on each tip

and the center section of the wing was reinforced with fiberglass. Servos were fitted into

each wing for the ailerons instead of having one servo in the center. This allowed the aile-

rons to also be used as flaps. Both the vertical and horizontal control surfaces were

enlarged proportionally with the wing. An O.S. .91FX engine powered the ATAs. Table

3.3 compares the final dimensions of each ATA. Notice that despite the significant

increase in size for the second plane, the weight actually went down. This was the result of

better planning and experience with building. The fuselage section for the ATAII was built

much more minimally, and yet strong enough to carry the avionics.

Table 3.3 Comparison of ATAI and ATAII Dimensions

ATAI ATAII ATAI ATAII

Wing Span 70 inches 81 inches Wing Loading 28.3 oz/ft2 19.7 oz/ft2

Wing Area 910 sq. in. 1053 sq. in. Horiz. Tail Area 166 sq. in. 193.2 sq. in.

Length 62 inches 69 inches Horiz. Tail Vol. 0.484 0.548


Table 3.3 Comparison of ATAI and ATAII Dimensions

ATAI ATAII ATAI ATAII

Weight 11 pounds 9 pounds Engine (hp) O.S. 61FX (1.9) O.S. 91FX (2.8)

3.5 ATA Testing

The ATA's were used to validate flight codes for autonomous flight. ATAI demonstrated

the ability to hold a bank angle commanded by the copilot. By varying the bank angle

command, the plane could fly circles under the computer's control. The plane was also

used in an attempt to test Phase two of reintegration. In the test, a mini van was driven

down the runway with the target light on top of it. The pilot attempted to maneuver the

plane into a position 10 meters behind the van so that the vision system could lock onto

the light and the plane could automatically hold its position relative to the van. Unfortu-

nately, ATAI was destroyed at one of these test flights during takeoff. Afterwards, it was

decided that this test was too risky and too difficult for the pilot to attempt with the ATAII.

The ATAII's role in PCUAV was primarily to test autonomous flight under GPS navi-

gation as would be required in Phase one. The plane did successfully fly a circle autono-

mously. It was in one of these circles that the plane was lost. The ATAII did not have any

pressure sensors, and relied on maintaining ground speed measured by GPS. When the

plane turned downwind its airspeed decreased to the point of a stall, the result of which

was described in 2.3.8. To fix this problem, NGMII and the Parent were fitted with pitot

tubes to measure airspeed, and all of the flight codes were modified to ensure that the

plane stayed above stall speed.

3.6 Chapter Summary

Chapter three described the designing, building, and testing of the PCUAV unmanned

aerial vehicles. These included an outboard horizontal stabilizer configured Parent vehi-

Section 3.5: ATA Testing 63

64 Chapter 3: UAV Building and Testing

cle, three versions of the Mini vehicle, and two avionics testbed aircraft. A detailed

description of the building process was given for those planes that the author worked most

on, namely the Parent, NGMII, and ATAII. The testing described included both structural

and flight testing. Details were given about how the results of the testing changed the

designs of the aircraft.

Chapter

4Structural Modeling of the Parent


This chapter lays out the work done on analyzing the structural dynamics of the OHS Par-

ent vehicle. The natural frequencies and mode shapes of the plane are discussed as well as

the process followed to achieve these results.

4.2 Structural and Inertial Properties of the Parent

Chapter three presented the building process and materials used for building the Parent. A

three view drawing of the plane appears in Appendix A. From these drawings and a

knowledge of the construction, an attempt was made to calculate the structural properties

of the plane. This section describes the process followed and the results obtained.

4.2.1 Area Moments of Inertia

The most important properties for the analysis were the area moments of inertia along

with the Young's modulus of elasticity. The moment of inertia of the tail booms are the

easiest to calculate on the aircraft. They are found from (eq. 4.1) and (eq. 4.2).

4 4

( = = r 2 (eq. 4.1)xx yy4

65

66

4 4n( r 1 - r 2 )

I = 2 (eq. 4.2)

Notice that JO is twice Ixx or I, or more appropriately, as for all symmetric shapes:

JO = xx + Iyy (eq. 4. 3)

The radii, axes, and origin for these equations are shown in Figure 4. 1.

Figure 4.1 Cross Section of the Parent's Tail Booms

The moments of inertia, IXX and Iyy, of the tail surfaces were estimated by only taking

into account the main spar and rear spar, which are both rectangular cross sections. These

moments of inertia were found about the area centroid, calculated by (eq. 4.4). X and Y

are the cartesian coordinates of the center of area of a cross section.

Y =

IAi

(eq. 4.4)

-i

X =

IA

Chapter 4: Structural Modeling of the Parent

The moments of inertia of a rectangle about its local axes are calculated from (eq. 4.5).

The moments of inertia of the two spars are combined, using the parallel axis theorem

given in (eq. 4.6) and (eq. 4.7). In these equations, the x'-y' axes are the local frame of ref-

erence of each rectangle, the x-y axes are the coordinates of the cross section at the area

centroid, and d is the perpendicular distance between the two axes. The length of the rect-

angle's side parallel to the x axis is denoted by b, while side h is parallel to the y axis, and

A is the area of the relevant rectangle.

bh 3b 3hIX= 2 IYY, = 2 (eq. 4.5)

- 2=xx IxIx, +Ad (eq. 4.6)

- 2I = +Ad (eq. 4. 7)

Finally, the wing's area moments of inertia were estimated as the total inertia of the

main and rear spars combined with the balsa and composite skin. The properties of the

spars were calculated with (eq. 4.5) through (eq. 4.7), while the properties of the skin were

computed using a spreadsheet, which appears in Appendix A. The skin was discretized

into lumped areas every 2.5% to 10% of the wing's chord and were added into the total

inertia of the wing through the parallel axis theorem. The contribution of the balsa and the

composites were calculated separately. After each component's contribution to the inertia

Section 4.2: Structural and Inertial Properties of the Parent 67

68

was calculated, they were multiplied by their respective Young's modulus and modulus of

rigidity, then added together to find the overall flexural properties of the wing.

Composite Layup Ba sa Skin

Foam Core

Balsa BalsaLeading Edge Main Spar Rear Spar Trailing Edge

Figure 4.2 Cross Section of the Parent's Wing

Figure 4.2 provides a cross sectional view of the wing near the wing root, revealing the

components used in calculating it's moment of inertia. The cross sectional properties

change along the span due to the spar ending and the addition of the diamond shaped sec-

tion of carbon fiber as seen in Figure 3.3. The flexural properties of the different compo-

nents of the vehicle are presented in Table 4.1. The wing is broken up into sections

denoted by inches from the wing root. Those sections not listed are calculated as a linear

function between the adjacent sections. For the wing and tail, the x-axis is parallel to the

Table 4.1 Cross Sectional Flexural Properties of Parent

Component EIxx (lbs in.2) EIyy (lbs in.2) GJo (lbs in.2)

Tail Booms 350,000 350,000 230,000

Tail Surface Roots 25,200 1,080,000 25,000

Tail Surface Tips 5,700 221,000 5,100

Wing (0-5) 1,327,000 6,720,000 160,000

Wing (5) 1,613,000 25,536,000 720,000

Wing (20-36) 4,872,000 50,400,000 1,800,000

Wing (51-90) 340,000 11,600,000 410,000


chord, while the y-axis is perpendicular to both the chord and span. The axes for the tail

boom are defined in Figure 4.1. Note that the fuselage was assumed to be infinitely rigid

for this analysis.

The properties calculated as described above were later validated by experiment. The

wing and tail sections were assembled and the main spar was placed in a vice. Deflections

were measured when weights varying from one to four pounds were placed at the wing tip

and on the tail root. The flexural properties were then calculated by (eq. 4.8) and (eq. 4.9).

In (eq. 4.8), F is a force down on the wing tip, L is the distance from the point of cantilever

to the point where the force was applied, and 8 is the deflection at that point. In (eq. 4.9),

T is the torque applied on the wing by the weight placed on the tail, L is the distance from

the vice to the point where the torque is applied to the wing, and 4 is the angle of twist, in

radians, of the wing due to the applied torque. This experiment provided EIxx and GJo for

the wing. In the next experiment, the tail boom was put in the vice near the wingtip attach-

ment and deflections were measured when weights were placed on the tail surfaces. Using

(eq. 4.8), EIxx and EIyy were found for the tail booms.

FL3 (eq. 4.8)36

G J - (eq. 4.9)

The values found from direct experimentation were used to modify those found from

calculation. The calculated values were all larger than those found from experiment. This

could be due to assumptions made about both the Young's modulus and the modulus of

rigidity of the different materials, as well as the fiber orientation of both the composites

and the wood. Imperfect lamination and fiber-to-resin ratios could also be factors. The val-

69Section 4.2: Structural and Inertial Properties of the Parent

70

ues listed in Table 4.1 are the final figures used during the analysis process. They were

further validated using Aswing, as described in Section 4.3.

4.2.2 Weight and Mass Moment of Inertia

Each component of the aircraft was weighed separately. The mass moment of inertia was

calculated for each as well, using the approximation that each resembled a rectangular

block. The mass moment of inertia of a block can be calculated by (eq. 4.10), where m is

the mass of the block, and a and b are the sides of the block perpendicular to the axis of

rotation. The moments of inertia of each component were summed around the center of

gravity of the plane using the parallel axis theorem given in (eq. 4.11), where R is the per-

pendicular distance between the local and global coordinate systems. The global coordi-

nate system has it's origin at the center of gravity of the aircraft. The positive y-axis

extends spanwise along the right wing, the x-axis runs backwards from the c.g., and, using

the right hand rule, the z-axis points straight up relative to the plane.

2 2

lxt m(a + b ) (eq. 4.10)Ix, =12

Ix Ix,+mR 2 (eq. 4.11)

Table 4.2 presents the weights and mass moments of inertia about the plane's c.g. for

each component of the aircraft and the total for the entire aircraft. These figures were later

used to check the computer model described in 4.3.

Table 4.2 Weights and Mass Moments of Inertia of Parent Components

Component Weight (lbs.) Ix (lb ft.2 ) ly (lb ft.2) Iz (lb ft.2)

Fuselage, Fuel, and Avionics 20.8 0.033 0.251 0.234

Wing 14.25 1.676 0.052 1.708

Right Boom and Tail 3.1 1.079 0.201 1.275


Table 4.2 Weights and Mass Moments of Inertia of Parent Components

Left Boom and Tail 3.1 1.079 0.201 1.275

Total 41.25 3.867 0.705 4.493

4.3 Analysis Process

The analysis of the Parent's structural dynamics was done using Aswing, which is a pro-

gram created by Professor Mark Drela that combines computational fluid dynamics with

structural finite element methods to analyze both the steady and unsteady aerodynamics of

flexible bodies. A description of Aswing appears in Appendix B. 1.

Code was written for Aswing to describe the dimensions, structural properties, and

aerodynamics of the Parent. This code can be found in Appendix B.2. In the computer

model, the aircraft is broken up into six components: the fuselage, wing, right tail boom,

left tail boom, right tail surfaces, and left tail surfaces. The fuselage was modeled as an

infinitely stiff, tapered cylinder with a radius of 6 inches at its thickest part. The engine,

avionics, and fuel tank were modeled as point masses attached to the fuselage. The wing

was modeled as a beam with elastic properties varying as described in Table 4.1. Aerody-

namic properties were assigned to the wing including the lift versus a curve slope, maxi-

mum and minimum coefficients of lift, and pitching moment coefficients. Ailerons are

modeled as changes in the lift distribution over the wing due to input deflection angles.

Each tail surface is modeled as a single beam element running from the tip of the vertical

stabilizer, down through the root, to the tip of the horizontal stabilizer. Like the wing, both

structural and aerodynamic properties are assigned to each tail cluster. Finally, the tail

booms are each modeled as cylinders with a diameter of 1.125 inches and the proper stiff-

ness properties. Each component is given a mass distribution of weight per unit span.

Once a model was complete, it was fed into Aswing where the geometry could be

viewed and validated in a plot. (See Figure 4.3) Aswing computes the weight of each com-

Section 4.3: Analysis Process 71

72 Chanter 4: Structural Modelinar of the Parent

ponent and the total weight of the aircraft, as well as the center of gravity. The model came

to within 1/4 of a pound of the actual aircraft weight, and the center of gravity within 1/4

of an inch. This was deemed as acceptable since the actual weight and c.g. change by a

larger amount during flight as the fuel level decreases. The mass moments of inertia of

some of the components are also displayed in Aswing, and were validated to be within 1%

of those calculated for Table 4.2.

Z

RIZ - -EM'EL - 2 ASWING 5.43

Figure 4.3 Aswing Geometry for Parent

Aswing allows the user to place test weights at various locations on the aircraft. With

this feature it is possible to measure the static deflections of the aircraft's structure when

weight is applied. It was easy to validate the model with the experiments described in Sec-

tion 4.2. The stiffnesses of the beam components in the model were modified until the

deflections measured were within 1% of those observed in reality.

The natural frequencies of the tail booms were observed during the experiments in

Section 4.2. The natural frequency of movement in the vertical direction was measured to

be 2.2 Hertz, while the movement in the horizontal direction occurred at 4.6 Hertz. These

72

were validated by Aswing, where the natural frequencies were computed to be 2.14 Hertz

and 4.35 Hertz respectively.

A velocity sweep was done for the computer model, ranging from 35 ft/sec, which is

about 2 ft/sec above stall speed, to 300 ft/sec. Figure 4.4 represents a sweep from 35 ft/sec

to 210 ft/sec. The text at the top of the figure displays sideslip angle and angle of attack in

degrees, airspeed in feet per second, coefficient of lift, coefficient of drag, Oswald effi-

ciency, and rotational accelerations of the aircraft in trim condition for each operating

point's velocity. The plot at the bottom of Figure 4.4 shows how the plane flexes for each

trim condition. The horizontal tail surfaces go from being bent up at low speeds to being

bent down at high speed. At low speed, the strength of the wingtip vortex is large, produc-

ing a large lifting force on the tails. As the speed increases, the vortex strength becomes

less, and the tails' lift is decreased. This can be seen in the second graph in Figure 4.4,

which plots the vortex strength along the span for each velocity (this is related to local

coefficient of lift). It is large at low speed, and goes to zero at high speed. At very high

speeds the negative pitching moment of the wings becomes larger than the positive pitch-

ing moment caused by the center of lift being forward of the c.g. The tail then needs to

push down to trim the aircraft. The tail switches from pushing up to down at around 90 ft/

sec. The top graph in Figure 4.4 plots the effective angle of attack, for each velocity, along

the span of each surface. The wing is displayed along the entire horizontal axis of the

graph, while the tail surfaces are overlaid near the center of the plot. The wing is twisted

by the moment applied to it by the tails. As the speed gets large, the tips of the wings are

twisted up. This results in the lift being concentrated at the tips as opposed to being spread

over the entire wing, as seen in the third graph, which plots the lift force distribution over

Section 4.3: Analysis Process 73

the span of the wing. In Section 4.4, the aircraft is analyzed at each of the velocities in the

sweep to determine how it's natural frequencies and mode shapes change with velocity.

" 0

0.000.000.000.00.D00.DD0.000.000.000.00

a"u

12.823.700.53

-D.39-1.D B-1.9 L9

-1.53-2. 12-2.40-2.70

V r35.050.070,90.0

110.0130.0150.0170. 0190.0210.0

CL1.0720.5260.2680.1620.1090.07810.050.0450.0360.030

CD

0.05050.00970.00320.00180.00 IL0.00130.00130.00140.00150.0017

e1.1521.1771.057D.BB3D. BDLD.3710.2080.1110.0500.035

0. 0

0.0D. 0

0. 0D. 0

D. 00.00. 0

a?0.0a. aa. a0.00.00.00.00. 00.00.0

- L

------------------------------------ ----------

1.5 --------. .. .. .. ---- --- ---- --- ---- --- .--- ---

2 [~/cV,

(CL)

0.50

1 . - -------- - - - --- ------- ----- -

flirt U5y

5-0. -- ---------- ------ ------------

S . 0 . .. .. .....

9/

.- -

------ -------------------------------

.. . . . .. . . .

-- -- - - -- - - -- - - ------- ---------.-- - - - - - - - - - -- - - . - - - - - - - -

----------- -------- - ------- ----- --- -

-- - .------------- --- ....... -----

sii

I

Figure 4.4 Velocity Sweep of Parent

8.00

Woff;L4. 0

0.0

-'4.0a

-8.01

.

74 Chapter 4: Structural Modelingz of the Parent

Section 4.4: Natural Frequencies and Mode Shapes of the Parent 75

4.4 Natural Frequencies and Mode Shapes of the Parent

The analysis described in the previous section provided information about the trim condi-

tions of the Parent for a variety of velocities. The aircraft was then analyzed at each of

these velocities to find the mode shapes and natural frequencies of deformation. When

these frequencies are plotted in the imaginary plane, they provide a root locus plot for the

aircraft (see Figure 4.5). The first 10 mode shapes and natural frequencies of the aircraft,

......... ............ .0

1 -0 1. . . .-...... 8.

4.0

2.020.0...............................

2.0

-202.0

-4.0

.................

Figure 4.5 Root Locus Plot for Parent

flying at cruise speed, are presented in Table 4.3. These are the first ten modes following


the six rigid body translational and rotational modes. All of these mode shapes appear as

figures in Appendix C. 1. Note that the natural frequencies of the first two modes stay rela-

tively constant and damping gets larger as speed increases to about 150 ft/sec at which

point the frequencies increase while the damping decreases.

Table 4.3 Mode Shapes and Natural Frequencies of Parent, Airspeed = 70 ft/sec, @ S.L.

All of these modes were found to be stable. Some have very small damping ratios.

These may not be realistic since Aswing does not include structural damping. The damp-

ing properties of the foam in the wing alone could be quite high, increasing the damping

Mode Number Mode Shape Natural DampingFrequency (Hz) Ratio

1 Asymmetric Tail Boom 3.145 .286Bending (x-z plane)

Asymmetric Tail Boom 4.113 .1162 Bending (x-y plane)

3 Symmetric Wing 5.395 .083Bending

4 Asymmetric Wing 5.511 .072Twist

5 Symmetric Wing 6.818 .058Twist

6 Asymmetric Horiz. Sta- 8.622 .091bilizer Bending

7 Symmetric Horiz 9.038 .086Stabilizer Bending

8 Second Wing 16.96 .049Bending

9 Wing Bending 18.01 .006(Fore-Aft)

10 Asymmetric Vert. Tail 22.29 .046Bending

7676

ratios in Table 4.3 by .02 or more. This would not significantly change the natural frequen-

cies, however.


Chapter four discussed the work done by the author in forming a computer model of the

Parent and using it for flight analysis. The calculations and experiments done to find the

plane's elastic behavior were presented. The basics of coding in Aswing as well as the

analysis procedure were discussed here and are elaborated on in Appendix B. Finally, the

mode shapes and natural frequencies found with Aswing were presented in Table 4.3.

Section 4.5: Chapter Overview 77

78 Chapter 4: Structural Modeling of the Parent

Chapter

5Aeroelasticity


Chapter four discussed the process of creating an aerodynamic and structural model of the

Parent using Aswing. Also included were results of studying the natural frequencies and

mode shapes of the flying aircraft. Chapter five continues to discuss the results of the

Aswing analysis. The aeroelastic properties inferred from these studies, including flutter

and divergence analyses are presented. Also included are servo aeroelastic properties of

the aircraft, obtained from these studies, that may be used to augment the current control

laws, used for autonomous flight of the Parent.

5.2 Aeroelasticity of the Parent UAV

As described in chapter four, the Aswing model of the Parent was run through a sweep of

velocities ranging from 35 ft/sec, which is just above stall speed, up to 300 ft/sec. Chapter

four mainly focused on the general trends of the mode shapes and natural frequencies at

speeds expected during PCUAV missions. Under any normal conditions the aircraft was

not observed to flutter or diverge. It was not until about 240 ft/sec that the first flutter

mode was encountered. This is recognizable on the root locus plot as the speed when the

root crosses to the positive real side of the imaginary axis. The first flutter mode observed

79

is the asymmetric tail boom bending mode, or mode one in Table 4.3. See Figure 5.1 or

Figure C. 1 for a visualization of this mode.

PCUAV PARENT Model 1.]Dp.Point.i 3Mad@ IJF - 3.CE UCI1cl/9C - 0.28028* - 0'

EL -AGUING 5.1.1

Figure 5.1 First flutter mode of OHS Parent

In this mode, as the wingtip pitches up, the inertia of the tail tends to twist the wingtips

more, causing more lift on the wing and a greater bending moment on the tail boom. At

slow speeds the tail also sees an angle of attack change when the wing tip twists up, and

the lift generated by the tail counteracts the inertia. As the speed increases however, the

inertia of the tails overcome the restoring lift, to the point where the mode becomes unsta-

ble. To test the effect of the tails' inertia on this flutter mode, the Aswing model was run

with simulated three pound weights added to each tail and a six pound counterweight

attached to the nose of the fuselage by an infinitely stiff pylon. The length of the pylon

was adjusted to maintain the original c.g. location. The same mode was observed to flutter,

but at a speed of 120 ft/sec. The symmetric wing bending mode, mode four of Table 4.3,

80

Section 5.2: Aeroelasticity of the Parent UAV 81

also encountered flutter at 160 ft/sec. These modes flutter in the same manner, with the

first mode being an asymmetric version of the second. See Figure 5.2 for a view of the

root locus plot when the weights were added to the tail.

90-1............... . ............................................

20.0 --+-+ + --. -+ +-+-+ -ir-.. -+- + -I/S .................. ................ ......... ....... ..... . Uyls'

2.0

D.0 a

1-L-

-2.19

-3.0

25.0 -20.0 -J.D -LG.a -5D 0.0 5.0 Da 1/s

Figure 5.2 Root Locus Plot of OHS Parent with Three Pound Weights on Each Tail andCounterweight Attached to Fuselage

This analysis amplifies the need for a stiff wing and light tail section when building

OHS aircraft. Another solution for flutter problems was found to be the addition of a coun-

terweight extending forward from the wingtip. This configuration was tested in Aswing,

with the same three pound weights on the tails, but with three pound weights attached to

each wingtip to balance the plane. With these counterweights in place no flutter was

82 Chapter 5: Aeroelasticity

observed, and in fact the modes became stiffer with increased airspeed. See Figure 5.3 for

the root locus plot with the counterweights on the wingtips.

0.0. ...........-- -.-.. - + -

-3.0

/ - r- cycles/s

2.0

10.

1.0

Aill: -

-Z. D

-20.0 .. .. ..... ... .... .. ..

-3D. 0

-20. -Z.0 1J.D 45a 0 ,,s-0

Figure 5.3 Root Locus Plot of OHS Parent with Three Pound Weights on Each Tail andCounterweight Attached to Wingtips' Leading Edges

5.3 Flight Dynamics of Parent

Figure 5.4 is a zoomed in version of Figure 4.5, to show in more detail the movement of

Secton .3:Fligt Dnamcs o Paent83

the phugoid, short period, dutch roll, and spiral modes of the aircraft.

a.

"-4-.-'

a.

Phugoid

a a a pitral

Dutch Roll

Short Period

2.0

-0.0

Frequency(Hz)

*-2.0t

Sigma = 0

Figure 5.4 Blow Up of Root Locus Near Origin

As expected, the Dutch roll mode becomes stiffer, the spiral mode becomes more sta-

ble, the phugoid becomes less stiff, and the short period becomes more heavily damped

with increased airspeed. These modes were also estimated using standard aircraft flight

dynamics equations. The frequency and damping ratio of the phugoid mode were calcu-

lated by (eq. 5.1), the short period by (eq. 5.2), the Dutch roll by (eq. 5.3), and the non-

oscillatory real root of the spiral mode by (eq. 5.4). See the list of symbols on page six for

definitions of the variables in these equations

(oph = U0

Z M

U0

phF2(L/D)

M +M.+ aq a u

sp 2osp

(eq. 5.1)

(eq. 5.2)

Section 5.3: Flight Dynamics of Parent 83

84 Chanter 5: Aeroelasticity

(1) -r+ 0, 1 Y +u~odr Y 1 r' dr = I ir( U1A (eq. 5.3)

u0 2o)dr( u0

LgN -LNXspiral = Lrp (eq. 5.4)

In Table 5.1, the results of these computations are compared with the results obtained

through Aswing for flight at cruise velocity and standard sea level temperature and pres-

sure. The phugoid modes have similar frequencies, but quite different damping ratios. The

reason for this comes from low drag calculated by Aswing, where the plane was modelled

as a much more clean aircraft with a L/D ratio of 35 instead of a more realistic number like

10. The short period frequencies appear to be quite different. However, the damping ratios

are quite high making the discrepancy less important. The difference is highly dependent

on c.g. location and center of pressure positions, which could be slightly different from the

actual aircraft in both models. Finally, the Dutch roll properties found by Aswing appear

to be consistent with those calculated by (eq. 5.3).

Table 5.1 Computed Flight Dynamic Modes Compared with Aswing Results

Mode Computed o Computed ( Aswing co Aswing (

Phugoid .07 Hz .22 .05 Hz .02

Short Period 2.5 Hz .78 .64 Hz .94

Dutch Roll .54 Hz .21 .41 Hz .20

5.4 Servo Aeroelasticity of Parent

Bode plots were created in Aswing to analyze the transfer functions between rates of rota-

tion and control input frequencies. These included roll rate versus aileron deflection, pitch

rate versus elevator deflection, and yaw rate versus rudder deflection. All of these bode

plots appear in full page format in Appendix C.2. The frequency range chosen for the


Section 5.4: Servo Aeroelasticity of Parent 85

Bode plots was from .01 Hertz to 18 Hertz. This is consistent with the control capabilities

of the flight avionics. Sanghyuk Park also produced similar bode plots using a rigid air-

craft model and aerodynamic properties found using a fluid dynamics program developed

by Prof. Drela called AVL. The details of these studies appear in [8]. These plots were

confirmed by running the Aswing model with near infinite stiffness, achieved by multiply-

ing all stiffnesses by 10,000. The bode plots for the two rigid aircraft models appear to be

nearly the same, with the gain obtained from Aswing being about twice as high as San-

ghyuk's model for the aileron and elevator plots.

Bode Plot of Roll Rate vs. Aileion Frequency forOHS Parent10 . 1. . ...............

............ igi...... I- -FlexibleI.... III)::::......~10~ ... ,...,. .

. . . .. .e e . . e . .. . . . e. . . . . . a . . . . e. . .. .. . . . . . . .

.................... . .. . .

.*11 . . . . . i e . . . N - I.. . . ........

110- I 1 0 1 0 10

Aileron Input Frequency (Hz);

.100...........................................

-2S'i e e t e ... .... a '

30010F 1o, 10C 160 101 o

A ileo mn I nput Frequ ency (Hz)

Figure 5.5 Bode Plots of Parent Roll Rate Response to Unit Aileron Input,Airspeed= 70 ft/sec. @ S.L.

From Figure 5.5, it appears that the flexible and rigid models are similar at low fre-

quencies. The Dutch Roll natural frequency can be seen as a resonance in the gain plots

between .4 and .5 Hertz. From this point on the gain drops off and the phase goes through

a 90 degree shift. The two plots differ above 10 Hertz however, where the flexible body

creates a resonance in the gain at around 17 Hertz, which is associated with the eighth

mode listed in Table 4.3.

Bode Pbt of Pitch Rate vs. Elevator Frequency for OHS Parent

....... .......

Flexible-...--. Aswing Rigid .

.. . .. .|.|. .|. . |.|.

10

02

0

10~

10Elevator input Frequency (Hz)

10Elevator In put Frequency (Hz):

1010 1

10 1

103

10310

Figure 5.6 Bode Plots of Parent Pitch Rate Response to Unit Elevator Input,Airspeed = 70 ft/sec., @ S.L.

The pitch rate versus elevator input bode plots are also similar at low frequencies for

rigid and flexible aircraft models. Both have a peak in the gain near the phugoid fre-

quency, although that of the flexible model is less damped. For the flexible plane another

resonance in the gain curve occurs at 5.4 Hertz, associated with the first wing bending

10'2

**101

100

~10,

10~

-100

-B-200

-3.CL _4M

~a-0010


.. . . .. . . . . .. ... . .. ... . . , , . ,

" ..... ..

mode, or mode 3 in Table 4.3. Lastly, near the upper frequency of the flexible aircraft,

there is a resonance at 18 Hertz associated with the fore-aft wing bending mode.

10 2

j

' 10CS

1010

300

200

100

0

-IL

Bode Pbt of Yaw Rate vs. Rudder Frequency forOHS Parent

-R igid

. . .. N .. .. .. . . . .. . . .... .. . . . --- Asw ing Rigd............................... .. ,,................. ...

- -- -- - ----- - ---- - --- --- -- - -- ---- - ---- -- - - - - -- -- --.k-- -

~7 10 1~ 10 10 10Rudder Input Frequency (Hz)

10~" 10~I 1~ 10* 10 10lR udder Input Frequency (Hz):

Figure 5.7 Bode Plots of Parent Yaw Rate Response to Unit Rudder Input,Airspeed = 70 ft/sec., @ S.L.

3)

10 :

Again, in Figure 5.7 the plots are similar, with a zero in the gain at .09 Hertz, and a res-

onance near the Dutch roll at about 0.5 Hertz. However, the flexible vehicle also produces

a significant gain peak at 5.5 Hertz, associated with the wing twisting mode 4 of Table 4.3.

In summary, the Bode plots for each of the three control axes are similar for both rigid

and flexible aircraft models at low frequencies. Both models provide information about

I:a.

.. . . . ... . . . . . . . . . . . . . ..

Section 5.4 Servo Aeroelasticity of Parent 87

88 Chanter 5: Aeroelasticitv

the flight dynamic modes. The flexible model does introduce information about how the

flexible modes found in Chapter 4 affect the controllability of the aircraft. This informa-

tion should be augmented into the control system of the Parent aircraft. The control system

could be modified to avoid exciting the control surfaces at the natural frequencies found in

the flexible aircraft, thus reducing the amplitude of the aircraft's deformations and conse-

quent fatigue, thereby increasing the longevity of the aircraft's structure. Servo-aeroelastic

instabilities will be avoided by decreasing the gain of the flight controller at the flexible

aircraft's resonant frequencies.

5.5 Chapter Summary

Chapter five continued the discussion started in chapter four of the results obtained from

Aswing, focusing on the aeroelastic behavior of the Parent. There was no flutter or diver-

gence observed at normal operating speeds, and the first flutter mode appears at over three

times the reintegration airspeed of the aircraft, this mode is the wing's first bending mode

seen in Figure 5.1. The flutter can be induced more quickly by adding inertia to the tail

surfaces. It was found that a solution for OHS flutter is to add counterweights to the lead-

ing edge of the wingtips, which eliminates flutter and causes the flutter modes to actually

become stiffer with airspeed.

The flight dynamics of the Parent were discussed including the phugoid, dutch roll,

short period, and spiral modes. Bode plots were created with Aswing for a flexible aircraft

model and were compared with plots created in Matlab with a rigid aircraft model. It was

found that the flexible modes of the Parent should not be ignored in the design of the con-

trol system of the aircraft since flexible modes are present within the bandwidth of the

flight controller.

88

Chapter

6Summary and Conclusions

6.1 Thesis Summary

This section summarizes the information contained in this thesis. Chapter one introduced

the concepts of PCUAV. Chapter two expounded on those concepts and presented the

methods and tools used to complete PCUAV's objectives. Chapter three discussed the

designing, building, and testing of the unmanned air vehicles used in PCUAV. Chapter

four described the work done by the author to model and evaluate the structural dynamics

of the Parent aircraft. Finally, Chapter five discussed the aeroelastic properties and flight

dynamics of the flexible Parent.

6.1.1 PCUAV System Summary

The PCUAV project was formed, under the MIT/Draper Technology Development Part-

nership, to fill a perceived national need in UAV technology. Current UAVs are either

large aircraft capable of flying long distances at high altitudes, or small, maneuverable air-

craft capable of up-close surveillance of a target near the point of departure. No known

UAV is capable of both long range and up-close surveillance. PCUAV is an attempt to

combine existing types of UAVs into a system of cooperative aircraft capable of up-close

surveillance at a distance.

89

90

To demonstrate the practicality of PCUAV, two types of UAVs were designed, built

and flown as a system of cooperating planes. The first aircraft, called the Parent, is repre-

sentative of a long range autonomous vehicle that can attain high altitude. The second air-

craft, called the Mini, is smaller, less detectable by an enemy, and more maneuverable.

During a mission, the Parent carries the Mini to a target site, releases it, provides a com-

munications link between the Mini and home base, and finally retrieves the Mini and

brings it home.

Chapter two of this thesis discussed in more detail the steps to the above mission, and

pointed out the key technologies that must be demonstrated for it to succeed. It was

pointed out that the most important of these keys is the reintegration of the two aircraft.

The chapter went on to describe the procedure followed (Phases one, two, and three), and

the tools used to accomplish demonstration, such as the outboard horizontal stabilizer Par-

ent aircraft, the three Mini aircraft, the Avionics Testbed Aircraft, and the avionics sys-

tems.

Chapter three brought to light the work done by the author and other members of the

PCUAV group in designing, building, and flying the three types of aircraft used in the

project. The advantages and disadvantages of each design were presented. Details were

given about the building methods utilized for each aircraft. The results of flight tests

achieved by the time of this writing were presented as well as plans for future flight tests.

6.1.2 Suggestions for UAV Improvements

This section provides information on what the author believes would be improvements on

the building techniques for these aircraft as described in this chapter. This is presented as

changes that would be advantageous if building a replacement for each of the aircraft.

Chapter 6: Summary and Conclusions

When building a replacement Parent, a big advantage would be gained by replacing

homemade steel wire landing gear with off the shelf composite gear. Not only would this

be easier to assemble, it would be stronger and more reliable. During flights with the cur-

rent plane the landing gear proved to be unreliable, causing landings to always be more

difficult and risky. In addition to changing the landing gear, the fuselage of the aircraft

could be modified to make the length of the nose gear shorter to reduce the moment

applied to it by the ground. Secondly, a gas powered two stroke engine should be consid-

ered for this plane as long as the spark ignition could be shielded so as not to interfere with

the flight computers. The power available, even from the Moki 2.10, is marginal when in

extreme flight configurations. Finally, the spars for the tail surfaces were made incorrectly

on the current plane. A future plane's tails should not have full thickness spars, but instead

have bass spar caps with balsa webbing between the ribs, similar to the design for the tail

of the NGMII.

Like the Parent, the NGMII's landing gear are inadequate. A future plane should uti-

lize heftier gear. The vertical tail surfaces should either be braced by flying wires or rein-

forced at the attachment to the horizontal stabilizer. The current tail feathers proved to be

fragile at that joint and had to be repaired repeatedly. The GM- 15 airfoil used for the plane

is quite thin and difficult to keep torsionally stiff even with full balsa sheeting. It is also

difficult to "build-up" due to the undercamber of the airfoil. It is probably worth a little

extra weight to build the wing with a foam core and fiberglass and carbon skins. The time

required would be significantly less and the quality better. Finally, it may be advantageous

to use 5/8 inch carbon tubes for the aircraft's fuselage. Dremel tools of this size are readily

available, but 3/8 inch tools were never found. These tools are necessary to shape the ends

of the carbon tube to interface with one another when building the truss. The author cus-

Section 6. 1: Thesis Summary 91

92 Chapter 6: Summary and Conclusions

tom made tools of appropriate size, however these broke easily and had to be remade after

forming just a few truss members.

The ATAII came close to perfection as a test aircraft. It was relatively easy to build and

flew well. The final demise of the aircraft resulted from the wing strength being inade-

quate for recovering from a dive while fully loaded. It would help to add either another

layer of fiberglass or carbon fiber to the wing root.

6.1.3 Flight Tests

At the time of this writing both the NGMII and the Parent had been flight tested for auton-

omous navigation and had independently flown their respective parts for phase one of

reintegration. The team was waiting for the weather to be good enough to test phase one

with both vehicles in the air. In this test, both planes will be taken off under pilot control,

the parent will be switched to computer control, and will enter its orbiting circle. The Mini

will then be switched to computer control at an arbitrary location and will fly a phase one

path to rendezvous with the Parent. It will follow the Parent at a distance of 40 meters. A

future flight test is planned in which phase one will be flown in the same manner, but the

vision system will take control and the aircraft will continue in a circle as they perform

phase two navigation until the two planes make contact. At the conclusion of these two

flight tests reintegration of two unmanned aerial vehicles will have been demonstrated,

fulfilling the primary goal of PCUAV.

6.1.4 Structural Modeling of OHS Aircraft Summary and Conclusions

Chapter four discussed the work done to put together a computer structural and aerody-

namic model of the OHS Parent vehicle. The structural properties of the aircraft's compo-

nents were derived and the process of creating the computer model in Aswing was

described. A description of Aswing appears in Appendix B as well. Chapter four went on

to discuss the analysis results pertaining to the natural frequencies and mode shapes of the

Parent aircraft. These mode shapes are all illustrated in Appendix C. All of the flexible

modes of the aircraft appear to be stable for normal flight speeds of 35 to 50 miles per

hour.

Chapter five continued a discussion of the analysis results. It was found that the Parent

flutters at about 230 ft/sec, well beyond speeds actually experienced in flight. This flutter

mode is the asymmetric vertical tail bending mode. It was found that this mode could be

made to flutter at slower speeds by adding mass to the tails, and could be eliminated by

adding mass to the leading edge of the wingtips. Chapter five went on to present the flight

dynamics of the Parent, comparing the results from Aswing with those calculated using

standard dynamics equations. Those flight dynamic modes also appear in the Bode plots at

the end of the chapter. These Bode plots describe roll rate versus aileron input frequency,

pitch rate versus elevator input frequency, and yaw rate versus rudder input frequency. A

comparison was made between the results obtained using Aswing's flexible model and

Sanghyuk Park's rigid aircraft model, as well as a rigid model created in Aswing by multi-

plying the stiffnesses of the plane by 10,000. From these plots it was found that some of

the flexible modes of the aircraft affected the effectiveness of each of the control surfaces,

particularly at higher input frequencies. It is recommended that this information be used to

modify the current flight code used for autonomous flight.

Section 6. 1: Thesis Summary 93

94 Chapter 6: Summary and Conclusions

References

[1] Beer, Ferdinand P., and E. Russel Johnston, Jr. Mechanics of Materials. 2nd ed. NewYork: McGraw-Hill, Inc., 1992.

[2] Drela, Mark. Aswing 5.4 Technical Description. Internet. March 1999. Available:http://raphael.mit.edu/aswing/

[3] Kentfield, J.A.C. Upwash Flowfields at the Tails of Aircraft With OutboardHorizontal Stabilizers. AIAA 98-0757, 36th Aerospace Sciences Meeting & Exhibit.January 12-15, 1998.

[4] Lennon, Andy. Basics of R/C Model Aircraft Design: Practical Techniques forBuilding Better Models. Ridgefield, CT: Air Age Inc., 1999.

[5] Mukherjee, Jason. Automatic Control of an OHS Aircraft. Doctorate of Philosophyin Department of Mechanical and Manufacturing Engineering, University ofCalgary, March, 2000.

[6] Nelson, Robert C. Flight Stability and Automatic Control. 2nd ed. Boston: WCBMcGraw-Hill, 1998.

[7] Omelchenko, Alexander. Communication and Video Surveillance in the Parent andChild Unmanned Air Vehicles. Master of Science in Aeronautics and Astronautics,Massachusetts Institiute of Technology. 2001.

[8] Park, Sanghyuk. Integration of Parent-Child Unmanned Air Vehicle Focusing onControl System Development. Master of Science in Aeronautics and Astronautics;Massachusetts Institute of Technology. 2001.

[9] Raymer, Daniel P. Aircraft Design: A Conceptual Approach. 3rd ed. Reston, VA:AIAA Education Series, 1999.

[10] Urbain Francois. Vehicle Design, Flight Control Avionics, and Flight Tests for theParent and Child Unmanned Air Vehicle. Master of Science in Aeronautics andAstronautics, Massachusetts Institute of Technology. 2001.

[11] van Schoor Marthinus C. and Andreas H. von Flotow. "Aeroelastic Characteristics ofa Highly Flexible Aircraft." Journal of Aircraft, Vol. 27, (October 1990): pg. 901.

95

References96

Appendix

Vehicle Drawings

97

98 Appendix A: Vehicle Drwns

A.1 Three View Drawings of PCUAV Parent Aircraft

a?

'.0

Figure A. 1 Orthogonal Views of OHS Parent

Annendix A: Vehicle Drawings

Section A.2: Three View Drawings of PCUAV NGM I 99

A.2 Three View Drawings of PCUAV NGM I

rt-9

du 43C ZD

Figure A.2 Orthogonal Views of New Generation Mini

99Section A.2: Three View Drawings of PCUAV NGM I

100

A.3 Three View Drawings of PCUAV NGM II

%QCU

Appendix A: Vehicle Drawings

CU

CU

C>C>

Figure A.3 Orthogonal Views of Second New Generation Mini

Section A.4: Three View Drawings of PCUAV ATA I&II 101

A.4 Three View Drawings of PCUAV ATA I&II

Figure A.4 Orthogonal Views of Two Avionics Testbed Aircraft

101Section A.4 Three View Drawings of PCUAV ATA I&II

PARENT CHILD UNMANNED AIR VEHICLES & DRAPER LABSNew Generation Mini II Drawn & Built by Jason KepterDesigned by: PCUAV Drawn Moxrch 29, 2002

1 Bult' FoQk 2001MASSACHUSETTS INSTITUTE OF TECHNOLOGY

Notes'D) Top Hatches Made Fron Two Layersof Fiberglass Velcroed to Fuselage.2) Horiz. Stab Should be Same Levelas Wing, w/o Incidence.3) L.E. Stabilizer is 24' Behind T.E. WingSee Attached 3 View for Configuration.

Made FromBolsa

Ply

IL j -Throttle Servo____-3/32' Balsa

3/8" Oak Block w/ 3/8 Bass Dowelfor Rubber Bands to Attach WingPlace for Proper C.G. _

7.

A

Front Bulkhead

Note' Make Holes In Forward Bulkheadfor Bottery Access. Mount Nose GearBetween the Holes, One Inch Leftof Center.

Firewall

Lock w/ 1/8'Screws <x2)

tap-I

CD

cIQ

~2

Ca

-a0'-1

z0

Ca

a

CD

0(Q0

It

)0 O

Section A.5: Building Plans for NGM II 103

Ai9

9 LA

1 I

x01

0

a-

IL

Ni -- o

Figure A.6 Building Plans for NGM II Wing and Tail

Xu Yu Ai* i lAixi real i Ai* i^29 Ai~xi^2yA . y y

0 0.0% 0.0% 0.00 0.001.25 1.3% 2.2% 0.27 0.46 0.003171 0.228438 0.132813 Front Spar A 0.000724 0.000421 0.036656 0.000165468 -6.92427 0.152032.5 2.5%1 3.0% 0.53 0.64 0.0032 0.546125 0.398438 0 0.001748 0.001275 0.354344 0.000954496 -6.65864 0.1418935 5.0% 4.1% 1.06 0.8e 0.005839 0.7565 0.796875 Front Spar x 0.004417 0.0046531 0.564719 0.003341479 -6.26021 0.228822

7.5 7.5%j 5.0% 1.59 1.s 0.005598 0.965813 1.328125 0.00 0.005406 0.007434 0.774031 0.005221435 -5.72896 0.1837210 10.0%] 5.6% 2.13 1.20 0.0055 1.125188 1.859375 Front Spar 1 0.006189 0.010227 0.933406 0.006963226 -5.19771 0.14858815 15.0% 6.6% 3.19 1.40 0.010827 1.3005 2.65625 0 0.014081 0.02876 1.108719 0.018311981 -4.40083 0.20969320 20.0%1 7.3% 4.25 1.54 0.010714 1.473688 3.71875 Rear Spar A 0.01579 0.039844 1.281906 0.023269061 -3.338331 0.11940625 25.0% 7.7% 5.31 1.63 0.010661 1.586313 4.78125 0.016911 0.050971 1.394531 0.026826353 -2.275831 0.05521630 30.0% 7.9% 6.38 1.67 0.0106341.652188 5.84375 Rear Spar xi 0.01757 0.062145' 1.460406 0.029028882 -1.21333 0.015656140 40.0%. 7.8% 8.50 1.66 0.021251 1.6661 7.43751 17.00 0.0354041 0.158052 1.474219 0.058982452 0.380418 0.003075150 50.0%1 7.2% 10.63 1.54 0.021283 1.598 9.5625 Rear Spar 1 0.034011 0.203521 1.406219, 0.054349104 250541810.13359860 60.0% 6.4% 12.75 1.35 0.021332 1.445 11.6875 0.0308251 0.249319 1.253219 0.044542003 4.630418 0.457377170 70.0% 5.2% 14.88 1.10 0.021397 1.226125 13.8125 0.026236! 0.295552! 1.034344 0.032168524 6.755418 0.97648680 80.0% 3.8% 17.00 0.80 0.021466 0.9488131 15.9375 ] 0.0203671 0.342117 0.757031 0.019324817 8.880418 1.69286190 90.0% 2.1% 19.13 0.44 0.021544 0.6194381 18.0625 0.0133451 0.389144 0.427656 0.008266602 11.00542 2.609427,95 95.0% 1.1% 20.19 0.24 0.010811 0.342125 19.65625' 0.0036991 0.212506 0.150344 0.001265438 12.59917 1.716149

100 100.0% 0.1% 21.25 0.03 0.01084 0.134938 20.718751 0.0014631 0.224583 -0.05684 0.000197369 13.66167 2.023116Lower face 0.0% 0.0% 0.001 000 0.2125

1.3% -1.7% 0.271 -o.35 0.004399 -0.17531 0.132813 -0.00077 0.000584 -0.367091 0.000135195 -6.92427 0.2109032.5% -2.3% 0.531 -0.48 0.002965 -0.4165 0.398438 1{-0.00123 0.001181 -0.60828 0.000514352 -6.65864 0.1314635.0% -3.0% 1.06 -0.64 0.00554 -0.561 0.796875 -0.00311 0.004415 -0.75278 0.001743663 -6.26021 0.2171277.5% -3.5% 1.59, -0.74 0.005398 -0.68744 1.328125 -0.00371] 0.007169 -0.87922 0.002550876 -5.72896 0.177163

10.0% -3.8% 2.13 -0.80 0.005348 -0.76606 1.859375 -0.00411 0.009944 -0.95784 0.003138555 -5.19771 0.14448615.0% -4.1% 3.19] -0.87 0.010651 -0.83406 2.65625 -0.008881 0.028292 -1.02584 0.007409477 -4.40083 0.20628120.0% -4.2% 4.25 -0.90 0.010629 -0.88506 3.71875 -0.00941 0.039525 -1.07684 0.008325754 -3.33833 0.1184525.0% -4.2% 5.31 -0.90 0.010625 -0.89781 4.78125 -0.00954 0.050801, -1.08959 0.008564482 -2.27583 0.05503130.0% -4.1% 6.38 -0.e 0.010627 -0.88613 5.84375 -0.00942 0.062102 -1.07791 0.008344605 -1.21333 0.01564540.0% -3.8% 8.50 -o.8i 0.021261 -0.8415 7.4375 -0.01789 0.158128 -1.03328 0.0150553 0.3804181 0.003077

-3.3%

-1.5%

10.6312.7514.88

-0.71 0.021272 -0.75863-o.59, 0.021286 -0.64813-0.451 0.0212911 -0.52063

17.00 -0.32 0.021293 -0.3867519.13 -0.17 0.021299 -0.246520.191 -0.1021.251

0.01065-0.031 0.010651

-0.13813-0.06481

9.562511.687513.812515.937518.0625

19.6562520.71875

_ _1 10.643754 _

Neutral AxisX= 7.057082Y= 0.191781

Moments of InertiaIx=ly=

0.42064321.6456522.0663

-0.01614 0.2034181 -0.95041 0.012242561 2.505418 0.13353-0.0138 0.2487771 -0.83991] 0.00894140414.6304181 0.456382

-0.01108 0.294079 -0.712411 0.005770881 6.755418-0.008241 0.3393651-0.00525 0.3847151-0.00147

-0.57853-0.43828

0.97162

Unit airfoil geometry

0.0031849841 8.880418 1.6792420.00129418 11.005421 2.579727

0.209331 -0.329911 0.0002031771 12.599171 1.690497-0.00069 0.2206750.12346 4.543024

Upper face x

50.0%,60.0% -2.8%170.0% -2.1%180.0%l90.0% -0.8%95.0% -0.5%J

100.0% -0.1%

airfoil wvith chord - 151

, ,

,

50

O

(+

0

-ol

9=

CA

00

i xi real xi

Appendix

BAswing and Related Code for theParent

B.1 Description of Aswing

Aswing is an analysis tool created at MIT by Professor Mark Drela. It combines computa-

tional fluid dynamics and finite element methods to provide analysis of aerodynamics,

structural dynamics, and control of flexible aircraft with high aspect ratio surfaces and

fuselage beams. The structural components of aircraft are modeled as fully nonlinear Ber-

noilli-Euler beams broken into finite elements. The fluid dynamics part of the program is a

lifting-line model that employs wind-aligned trailing vorticity, a Prandtl-Glauert com-

pressibility transformation, and local-stall lift coefficients to predict flight characteristics.

It is possible to predict divergence and flutter speeds, aileron reversal speeds, the deforma-

tion of the aircraft and its effects on stability and control, and the stresses in the structural

components.

The code written for the model of the Parent appears in the next section. The first sec-

tion of the code describes the units used for the model, in this case, the length unit is .0833

feet or one inch, time units are in seconds, and force units are in pounds. The constant

block defines gravity in inches per second squared, sea level density in slugs per square

105

inch, and the speed of sound in inches per second. The next section, the reference block,

lays out the wing area, chord, and span in inches. The point where global velocity, acceler-

ation, and momentum are calculated is defined in this section as well. In the case of the

parent, the global coordinate system has its origin at the firewall and at the same vertical

level as the bottom of the wing. The velocity measuring point defined in the reference sec-

tion is 21.4 inches behind the origin. This is roughly the center of gravity of the aircraft.

The weight block of the code defines point masses attached to the structural beams of the

aircraft. Weights and their placements are defined for the engine, the avionics in the fuse-

lage, the avionics in the wingtips, and the fuel tank. Two weights were defined for the fuel,

one for a full tank and the other for half of a tank. Aswing ignores a line when a "!" is

placed in front of it, so the plane can be flown with either fuel condition just by moving

the "!." The last weights defined are test weights. These are weights that are placed either

on the tail or the wingtip to simulate actual experiment, as discussed in Section 4.2. The

engine thrust information is given in the engine block. The joint block defines the points

on each structural beam's local axis where they are attached to another beam. The fuselage

is labeled beam 0, the wing is beam 1, the right tail boom and tail surface are beams 2 and

3, while the left boom and tail surface are labeled -2 and -3. The last block before the

geometry definition is the ground block, which defines where the plane is held while

doing static tests. In this case, it is right behind the spar, which is where the wing was held

during testing.

Each geometry element is defined as either a surface beam or a fuselage beam. Both

types of beams are given geometry definition as well as mass and stiffness distributions.

Surface beams are also given aerodynamic properties, as described in Chapter 4.

106 Appendix B: Aswing and Related Code for the Parent

Secton .2: swig Coe fr th Paent107

B.2 Aswing Code for the Parent

NamePCUAV PARENT Model 1.1End

UnitsL 0.083333 ftT 1.0 sF 1.0 lbEnd

Constant#g rhoSL VsoSL386.16 1.1468E-07 13380.0

End

Reference# Sref Cref Bref

3825 21.25 180

# Xmom Ymom Zmom21.4 0.0 0.0

# Xacc Yacc Zacc21.4 0.0 0.0

# Xvel Yvel Zvel21.4 0.0 0.0

End

Weight# Nbeam t Xo Yo Zo Weight CDA* 1.0 1.0 1.0 1.0 1.0 1.0 1.0# Engine

0 0.0 -3.0 0.0 -2.0 3.0# Payload

0 12.0 12.0 0.0 -5.0 5.0# Fuel

0 9.0 9.0 0.0 -5.0 3.00 0 9.0 9.0 0.0 -5.0 1.5

# Wingtips1 85.0 12.0 85.0 3.1 0.51 -85.0 12.0 -85.0 3.1 0.5

# test mass

Vol

0.1 0.0

0.0 0.0

0.0 0.00.0 0.0

0.0 0.00.0 0.0

83. 83. 85.0 7.085. 18. 85. 3.185. 83. 85. 7.

3.0 0. 0.4. 0. 0.3. 0. 0.

Engine# Keng Nbeam t Xo Yo Zo Fx/Peng Fy/Peng Fz/Peng* 1.0 1.0 1.0 1.0 1.0 1.0 1.0

1 0 0.0 -3.0 0.0 -2.0 20.0 0.0 0.0End

E3n

End

Section B.2: Aswing Code for the Parent 107


Joint# Nbeaml Nbeam2 ti t2

0 1 18.0 0.01 3 85.0 20.01 -3 -85.0 20.03 2 82.0 0.0

-3 -2 82.0 0.0End

Ground# Nbeam t Kground

0 18.0 0End

Beam 0Fuselage

t x y z radius mg Cdf Cdp+ 0.0 0.0 0.0 -6.0 0.0 0.0 0.0 0.0* 1.0 1.0 1.0 1.0 1.0 0.15 0.005 0.3

0.0 0.0 0. 0. 2.0 2.0 1.0 1.09.0 9.0 0. 0. 6.0 2.0 1.0 1.09.0 9.0 0. 0. 6.0 2.0 1.0 1.0

29.0 29.0 0. 0. 6.0 1.0 1.0 1.029.0 29.0 0. 0. 6.0 1.0 1.0 1.048.0 48.0 0. 0. 0.0 1.0 1.0 1.0

End

Beam 1Wing

t chord x y z Xax twist Ccg mgnn*1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.00.0 21.25 18.0 0.0 0.0 0.39 0.0 0.25 2.3890.0 21.25 18.0 90.0 3.1 0.39 0.0 0.25 2.38

t alpha Cm CLmax CLmin*1.0 1.0 1.0 1.0 1.00.0 2.2 -0.05 1.1 -0.8

90.0 2.2 -0.05 1.1 -0.8

t dCLdadCLdF1 dCMdF1 dCLdF2 dCMdF2*1.0 1.0 1.0 1.0 1.0 1.0-90.0 6.28 0.0 0.0 0.0 0.0-78.0 6.28 0.0 0.0 0.0 0.0-78.0 6.28 0.055 -0.011 0.055 -0.011-36.0 6.28 0.055 -0.011 0.055 -0.011-36.0 6.28 0.0 0.0 0.0 0.00.0 6.28 0.0 0.0 0.0 0.036.0 6.28 0.0 0.0 0.0 0.036.0 6.28 -0.055 0.011 0.055 -0.01178.0 6.28 -0.055 0.011 0.055 -0.01178.0 6.28 0.0 0.0 0.0 0.090.0 6.28 0.0 0.0 0.0 0.0

t mg Cea Nea GJ Elnn Elcc* 1.0 0.06 1.0 1.0 10000.0 168000.0 168000.0

0.0 1.0 -1.5 0.35 16.0 40.0 7.95.0 1.0 -1.5 0.35 16.0 40.0 7.95.0 2.0 -0.1 0.32 72.0 152.0 9.6



20.020.036.036.051.051.090.0

2.02.02.01.01.01.01.0

0.00.00.03.63.33.33.3

0.320.320.320.290.280.280.28

End

Beam 2Right Stabilizer Unit

t x y z chord X0. 82.0 85.0 7.1 0.01.0 1.0 1.0 1.0 1.0

-17.75 0.0 0.0 17.75 9.00.0 0.0 0.0 0.0 15.00.0 0.0 0.0 0.0 15.0

30.75 0.0 30.75 4.25 9.0

t dCLda dCLdF3-17.75 5.28 0.055

0.0 5.28 0.0550.0 5.28 0.0

30.75 5.28 0.0

t* 1.0-17.75

0.00.0

30.75End

Cea1.0-2.36

-6.64-1.69

2.68

ax twist a0.0 0.0

1.0 1.00.56 0.0

0.73 0.00.4 -8.0

0.0 -8.0

lpha0.0

1.00.0

0.00.0

0.0

dCMdF3 dCLdF4 dCMdF4-0.011 0.0 0.0

-0.011 0.0 0.00.0 0.055 -0.011

0.0 0.055 -0.011

Ccg mg1.0 1.0

-1.44-4.95 00.0 0.3.6 0.

mgnn1.0

0.02 0..04 1.004 1.002 0.1

119

GJ Elnn Elcc0.0 168.0 168.0

510.5 1313.9 33.782547.2 6457.4 149.4

2547.2 6457.4 149.4510.5 1313.9 33.78

Beam -2Left Stabilizer Unit

t x y z chord Xax twist alpha+ 0. 82.0 -85.0 7.1 0.0 0.0 0.0 0.0* 1.0 1.0 -1.0 1.0 1.0 1.0 -1.0 -1.0-17.75 0.0 0.0 17.75 9.0 0.56 0.0 0.00.0 0.0 0.0 0.0 15.0 0.73 0.0 0.00.0 0.0 0.0 0.0 1 5.0 0.4 -8.0 0.030.75 0.0 30.75 4.25 9.0 0.0 -8.0 0.0

t dCLda* 1.0 1.0-17.75 5.28

0.0 5.280.0 5.28

30.75 5.28

t* 1.0-17.75

0.00.0

dCLdF3 dCMdF3 dCLdF4 dCMdF41.0 1.0 -1.0 -1.0

0.055 -0.011 0.0 0.00.055 -0.011 0.0 0.00.0 0.0 0.055 -0.0110.0 0.0 0.055 -0.011

Cea Ccg mg mgnn GJ Elnn EIcc1.0 1.0 1.0 1.0 10.0 168.0 168.0-2.36 -1.44 0.02 0.19 510.5 1313.9 33.78

-6.64 -4.95 0.04 1.0 2547.2 6457.4 149.4-1.69 0.0 0.04 1.0 2547.2 6457.4 149.4

180.0180.0180.0170.041.041.041.0

300.0300.0300.069.069.069.069.0

29.029.029.021.0

2.02.02.0

+ -*



30.75 2.68 3.6 0.02 0.19 510.5 1313.9 33.78End

Beam 3Right Boom

t x y z radius Cdf Cdp+ 0.0 0.0 85 3.1 0.5625 0.0 0.0* 1.0 1.0 1.0 1.0 1.0 0.005 0.3

20.0 20.0 0.0 0.0 0.0 1.0 1.083.0 83.0 0.0 4.0 0.0 1.0 1.0

t mg mgnn mgcc GJ* 1.0 0.016 0.001 0.001 23.0

20.0 1.0 9.4 9.4 10000.083.0 1.0 9.4 9.4 10000.0

End

Elnn35.0

10000.10000.

Elcc35.0

0 10000.00 10000.0

Beam -3Left Boom

t x y z radius+ 0.0 0.0 -85 3.1 0.* 1.0 1.0 1.0 1.0 1.0

20.083.0

20.0 0.0 0.0 0.083.0 0.0 4.0 0.0

t mg mgnn mgcc* 1.0 0.016 0.001 0.01

20.0 1.0 9.4 9.483.0 1.0 9.4 9.4

End

Cdf Cdp625 0.0 0.0

0.005 0.31.0 1.01.0 1.0

GJ Enn Elcc23.0 35.0 35.0

10000.0 10000.0 10000.010000.0 10000.0 10000.0


5

Appendix

Aswing Results

113

Appendix C: Aswing Rest

C.1 Mode Shapes

m

In

Lb

Id,

-I

-~1

A,-o0

HzLu

cr0~

C-,

4-ICCL

uS )

V Co

E- %J

1<

N

eIn 4n

I Cri

I II

Figure C. 1 First Mode Shape of OHS Parent, Asymmetric Vertical Tail Boom Bending,Airspeed = 70 ft/sec, @ Sea Level

114ADDendix C: Aswing Results

Section C. 1: Mode Shapes

m

InU,

LO3r

-I

-IJ

CL

CL.

C

C3

U 3

r- r'.i

-4 -- 1.

(n

- E L. I

1<

N

en* Lfl

h4 -J= LIi

Figure C.2 Second Mode Shape of OHS Parent, Asymmetric Horizontal Tail BoomBending, Airspeed = 70 ft/sec, @ Sea Level

1 15115

Appendix C: Aswing Results

m

In

In

h4

0

I Ni

1,- -

Figure C.3 Third Mode Shape of OHS Parent, Symmetric Wing Bending,Airspeed = 70 ft/sec, @ Sea Level

116

-D

I-

CD

'-I

SrU') CU

Mn 00to M II3I

-- C&; O%

Crn

C

0

a,

13-

Co,

4aC

("4

M

U5C

U-i 6

N -

Figure C.4 Fourth Mode Shape of OHS Parent, Asymmetric Wing Twist,Airspeed = 70 ft/sec, @ Sea Level

Section C. 1: Mode Shapes 1 17

M

Ln

cD

Ln

0U.,.0n) U2

I E

I= II

117


Ln

LOU,

:3LnE

'I I r2

Figure C.5 Fifth Mode Shape of OHS Parent, Symmetric Wing Twist,Airspeed = 70 ft/sec, @ Sea Level

118

F-

:z

a_-

q:

0L-

m c

0,do C

C4.'C0.

0.0

Section C. 1: Mode Shanes 1 19

rn

in

CC1

--

CL.

CDU

U r

U0 LflCD ED

C3

4- onc

D. MD.- 0 1, 11 I

Mo3 2: Ce. %J 4b.

4<

N

01171nI LFlI CU

1 11

= Lii

Figure C.6 Sixth Mode Shape of OHS Parent, Asymmetric Horizontal Stabilizer Bending,Airspeed = 70 ft/sec, @ Sea Level

119

Appendix C: Aswing Rest

Ln

Lb

1< 0r o

III2

Figure C.7 Seventh Mode Shape of OHS Parent, Symmetric Horizontal StabilizerBending, Airspeed = 70 ft/sec, @ Sea Level

120

F-

LU

-

ED0)

M

~r c

I- II ICin to.

C-C0

CyJ

Atmendix C: Aswing Results

-4

C3 w

w o'

a m C

CL a

Figure C.8 Eighth Mode Shape of OHS Parent, Second Wing Bending,Airspeed = 70 ft/see, @ Sea Level

Section C. 1: Mode Shapes 121

m.1

LO

-4

-4ID

DrH-

ulcr0~

In U3F)Lf

Appendix C: Aswirng Rest

z

Lfn

4<

N

0W7 0I riJ

1 11

Figure C.9 Ninth Mode Shape of OHS Parent, Fore-Aft Wing Bending,Airspeed = 70 ft/sec, @ Sea Level

122

-4

a_--D

m

CL

a,

C -

a.- C I I I

Annendix C: Aswing Results

Section C. 1: Mode Shapes

m

LA

LD

V)cc

r- Da-Z

-3

a- 03 1't

C3 X: L- J *& PLU

Figure C. 10 Tenth Mode Shape of OHS Parent, Asymmetric Vertical Tail Bending,Airspeed = 70 ft/sec, @ Sea Level

123


In

cc

L0P-0 Lfl

Figure C. 11 Eleventh Mode Shape of OHS Parent, Symmetric Vertical Tail Bending,Airspeed = 70 ft/sec, @ Sea Level

124

-D

CL

0)

CU

I

Li-

C4

Z-IL

:F'

La

o aC3 C3

Section C.2: Bode Plots 125

C.2 Bode Plots

PCURV PARENT Model 1.1

Operating point: 3

Response to unit 6F,

i.D

n ]

ref (2 = 1.00000 deg/si n

S . . .. L ... . . . . .. .. . .. .. .. .

.. . .... ... . ... . .- -.. . . . --.-. . . . .

.L .t J . .J LJ . . . . . . . . ............. .... . . . . L J --.--.. ....- - -- - . J . J. J-

0.01 0.1 1.0 10.0 cycles/5 iaao

Figure C.12 Gain Plot of Roll Rate vs. Aileron Input Frequency for Flexible OHS Parent,Airspeed = 70 ft/sec, @ Sea Level

.I


126 Annendix C: Aswin2 Results

PCURV PARENT Hodel 1.1

Operating point: 3


270.0

210.0-

90.00.

I 0K deg............. ..............

e e . e . . . . . . -- . .. .j 1 1- . . --------. = . . e e e e e

0.1 1.0 10.0 cycles/s 100.001

Figure C. 13 Phase Plot of Roll Rate vs. Aileron Input Frequency For Flexible OHS Parent,Airspeed = 70 ft/sec, @ Sea Level

K

126



Operating point: 3

Response to unit 6F.

ref Qh = 1.00000 deg/s

.................................................................... a............................... .,..... ........... ..... . . . e ... . . a . ._ e ... .d 1 r1 . . . . .. . e . . sr . . .. | -- -ri~ -N -v i %r--

r- 1-1. . . . . . a . a . . . . . . . . . . . . . a . . .I e . . .,r- ---

.......... ........ ,............. ,... ,......... e.,......... ,.....,........ ,...,............................,.. ... .... ..... , --.----. .----. .-

. . . . . . eer -ri. . ... r.. . . . s . . . . . . . . . . . . . . . .,- --- 4 ---

S . .. - . . .j_.. ..... L .... J L A L. . J. . . . . .e . . J . . . . . . . ._ JJ J. . . . .%

s.... . . ..... . . . . . . . .

- I- . . .. . . - . r - ,- * .F .r . " 1 F. 1*. -- - -- 4

0.. . .. . . . . . . . . . . . e

. . . . e . . , , . . . . e .e . . . r a r . . . . e -- -- . . . . . . . . .

e~ , , , e . .. . . . .. ,. .e l . . . . a . e . . . . .

a.. .. . . . .. .. , , e,., , , , . .. , , , , , , , , e . , , ,

.... ....... ... .. . .L. ........L...... ......L.L.JJ.................. .L JJ ............--.. .. ... J- . ,

..... . . . . . . . . . . . .. . r . . . . . . . 14 - - - - - - - -

. ........I . ...... .1 .0 .. ... ,-e/ 1 0 - 0 -

F , C 1 G Plot . . ..R a t v E l v t .I F uen c y F ex i e O H S are..... A Irsd. . . . Se - L " ... a . . .Leve l . . ... -

S a e a . a. . . . . . . . . . L .. . . a a a e a a a e. . . . . . . .

Figure~~~ C. 14 Gai Plo of Pic ateae vs .lvao Inu Freqeny Flxil OH Parent.. ..

aeAir esee = 70 f/ec @ Sa eeleee. eeeeea

128 Appendix C: Aswing Results

PCURV PARENT Hodel 1.1

Operating point: 3

Response to unit 6F.

L Og deg

-BD.-

. .--- ..... .... ,.. J . . . . L. .. ........ L.. -.L.-J.---A.- .

D D 0 . .. ,. . .. .. . L . -. . .... . . .---. . .. . .... ..

Tt

-540.00.01 0.1 1.0 10.0 cycles/s in0o

Figure C. 15 Phase Plot of Pitch Rate vs. Elevator Input Frequency, Flexible OHS Parent,Airspeed = 70 ft/sec, @ Sea Level



Operating point: B

Response to unit 6 F.

ref 2z = 1.00000 deg/s

....... A.L... ............. .....

10t L J. L.LJ. .. . . .JtJ. . L.J.J.J.L.

D.]

2- I 1- 4 1~ ~~~- - --- --4 -- 4 --- 44-2 - 4 4 4 44

0.01 0.1 1.0 10.- cycles/s 100.0

Figure C. 16 Gain Plot of Yaw rate vs Rudder Input Frequency For Flexible OHS Parent,

Airspeed = 70 ft/see, @ Sea Level



Operating point: 3


£ Oz deg

60.0 ......

. A .. . . .. .... .. ... . .. L... ... -4 . . . . ..... .. . ... .L . . . .. . . .. . . .. . . .

1 0 . . . . . .. . . . . .. . . . . . . . . . ... ...

......... I.... I... -.-.. -. J -6 ........ L...- L...- .J.. L.JL ........ L....L..J..J..L LJ. LL........ L.. . . J-..J.A - L J

-3DD.+0.01 0.1 1.0 10.0 Cycles/s 100.0

Figure C. 17 Phase Plot of Yaw Rate vs. Rudder Input Frequency For Flexible OHS Parent,Airspeed = 70 ft/sec, @ Sea Level


PCURV PRRENT Model 1.1

Operating point: 3

Response to unit 6FI

ref = 1.J000 deg/s

,.,..,,. .. .. ......

........................................ I .I........ ............ ...........................

.... ......................... ... I

- L.--------. ..-. .. ... .L L---------~t~ .

D.]0.01 0.1 1.0 ID. Ucycles/s -0

Figure C. 18 Gain Plot of Roll Rate vs. Aileron Input Frequency for Rigid OHS ParentAirspeed= 70 ft/sec, @ Sea Level



Operating point: 3

Response to unit 6Fj

L 0 deg

90.00.01 .1 1.0 10.0 cycles/s 10.0

Figure C. 19 Phase Plot of Roll Rate vs. Aileron Input Frequency For Rigid OHS Parent,Airspeed = 70 ft/sec, @ Sea Level



Operating point: 3


ref Qi = 1.DOOOOdeg/s100.0"I ......-----.

- .... L... .J..t. J........L....J...L...L.L . . J....J...L...J.J. J.L -. . .. J J. 6- - -- + -*4 ..... . .i-.4.+. . . . . ..- -- --- - ..--..-- I

-- -- -- - -4-------- -- - - - --- I-

.... .......... ... 6.....................

................ .... ... L.J4.....--..j-1JJ.

1.0.

1~~ DL ..... - .. ......_. J . .. tL .L.. . J.-- -- ... t..tJ.J . . -.1....I. .I 1.- .-

. 2-~ ~ -1- 4- -- -- 14 -4 -44

D.]0.01 0.1 1.0 10.0 cycles/s 100.0

Figure C.20 Gain Plot of Pitch Rate vs. Elevator Input Frequency, Rigid OHS Parent,Airspeed = 70 ft/sec, @ Sea Level



Operating point: 3

Response to unit 6F4

L Ow deg

s . .. .. . . ... . ..... . ... . . .... .. . . .... . ......

-30D.0 1 11:

0.01 0.1 1.0 10.0 cycles/s [00.0

Figure C.21 Phase Plot of Pitch Rate vs. Elevator Input Frequency, Rigid OHS Parent,Airspeed = 70 ft/sec, @ Sea Level



Operating point: 3

Response to unit 6 F3

ref Q, = 1. DOOtJ deg/s. . . . . . . . . . . .. .. . . . . . . . - - - - - - - --. . . . . . . . .

.. . C.. . . . ... . . 1. . . .... . . . ..- . . . ...-. .-. e . .- -.. . . ....- -. . . . ....

. . . .. . . e . . . . . . ... . . e . . . . .,e - - - -i-, . . .. . .... . . .. ........

I-- -L--I--I- A-.LJ. -..... . ... L Le . J."J - I. J L .. J .. - . . ...

"I-- - - - - -------- - -e------- -- - - -

. . . . . . . . - . . - - -

.. . . . .. -A A , -, -. . A -

F,... .. ....4. ... . . . ... .F. 4. 4 .. 4 4 -.- . . .... .... . . .4 4a~~~~~~~~~ ~~~~ . . . . .

. a . . . s . ..................... . . .. . .. .... . e. ... ,

D. ...-

S . . . . . . .... ... ... .. -----.. --- . . . . . . . . .. . . ... . . . . . e - .L . e1- -1 --------. . . . . . a. . . . .. .... ... . . - . . -, -

. . ...... .... . ... L . . . . . . .

, ~ ~ ~~~ -. . .. . .... ............. a e.. . . . e a , e e . .

e- 4 . - F -. . . . . . . . . F F 4 4 .-0 l- . . 4 4 4 . . .

0.01 0.11.0 10.0cycles/s 10.

Figure C.22 Gain Plot of Yaw rate vs. Rudder Input Frequency For Rigid OHS Parent,

Airspeed = 70 ft/sec, @ Sea Level



Operating point: 3

Response to unit 6F3

L Oz deg

. .L - . . . . . . .- . . - - . . . . - . L - . . J. L . . . . . . . L L - . .-.

D . .. .. . e- - r . .- . .- -r .r . T ,--- - re -r .I -f i r . -- -- - r ... .- .1 .1 1" , -- -- - , ,e ,- e- .,

1 .. ,.... . ,,, , ......... ..........

-120. 00.01 0 1 1 0 , a cycles/ , 1000

Figure C.23 Phase Plot of Yaw Rate vs. Rudder Input Frequency For Rigid OHS Parent,Airspeed = 70 ft/sec, @ Sea Level

. a . . a e , e . . , a . . s . . , a . . a , 'e .', a e a a.a . . a aa a , aa. , .,1 , e'. e a ea .. . . H K.. ) ', w 'k 1 e, , e a,

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