senior project final report

GUst Alleviation and Controls - Senior Project Report System Characterization and Identification of a Blended Wing

Body Aircraft

Submitted By: Tuan Dinh Jr (Team Lead), Dwight Nava (Co-lead), Reginald

Guinto, George Paguio, Tanner Clark, Jason Kong, Dong Jin Ryoo, Bill Wogahn,

Arya Williams, Crystal Nunez, and Anahi Hernandez

Project Advisor: Professor Steven Dobbs

California Polytechnic University, Pomona Aerospace Engineering Department

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Executive Summary

The GUst Alleviation and Controls (GUAC) research project is a continuing multi-year

project. The idea behind the project is to research and revolve around the idea of a gust

alleviation system as well as aero-elastic flutter. The project used an existing aircraft, X-56 Dart,

which is a high aspect ratio blended wing body design modeled after Lockheed Martin/ NASA

X-56 MUTT. This year the primary objectives were modifying the blended wing body wind

tunnel model fabricated by the 2013-2014 FALCON Club senior project team with the addition

of a horizontal tail to add more stability to the aircraft, obtain trim flight in the Cal Poly Pomona

Low Speed Wind Tunnel, gather test data for aircrafts with high aspect ratios for short period

mode stability, measure the model’s gust response and create a stability augmentation system

alleviating the gust response.

In achieving set goals and objectives, Team GUAC had created an approach in

completing the research project. This year the GUAC had approximated the X-56 type model’s

CG and AC locations. This data helped lead to the design of a horizontal tail with functioning

elevator. With addition to the new horizontal tail, the gimbal mount were adjusted to

accommodate the length of the tail to avoid any interference. With these additions to the existing

model, Team GUAC was able to obtain trimmed flight.

After obtaining trimmed flight, Team GUAC had created three different tests to obtain

data for the short period mode stability as well as gust response. The first test performed was the

elevator pitched excitation. This was use to simulate a pitch doublet maneuver that current test

pilots perform during flight test. The second test is the stick hit excitation test, which is

performed by having a long and slender dowel tap the nose boom of the model after having

obtain trimmed flight. This simulates a disturbance that the nose will encounter during gusts.

Lastly, the third test was the use of the wind tunnel dual gust vane excitation system within the

subsonic tunnel to create sinusoidal disturbances in the airflow. This test was used in order to

find the model’s natural frequency and response to the artificial gust.

The data collected from both the stick hit and the gust response test helped created three

important graphs. The three graphs are the pitch rate frequency vs velocity (Figure EC-1) with

theoretical vs. experimental correlation, pitch rate damping vs velocity (Figure EC-2) with

theoretical vs. experimental correlation, and pitch rate coefficient vs gust frequency (Figure EC-3

).These graphs help showcase the stability of the model as well as its response to gust. This data

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will then lead to developing the stability augmentation system. Below are the three important

graphs taken from this year’s project:

Figure EC-3 Model Maximum Peak to Peak Oscillating Pitch Amplitude vs. Gust Vane Frequency @ V = 90 ft/s

Team GUAC has added on to the multi-year project and is looking forward with new

ideas and technique in achieving a stability augmentation system to respond to disturbances such

as gust. This year’s research wouldn’t have been accomplished without the guidance of Professor

Steven Dobbs as well as the faculty, staff, and other colleagues. Team GUAC has provided

further foundation for the next team to continue on the research in gust response alleviation for a

high aspect ratio UAVs and other aircraft.

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Damping vs. Velocity

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Figure EC-1 Frequency Vs. Velocity Figure EC-2 Damping Vs. Velocity

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Table of Contents

1.0 Introduction ........................................................................................................................ 12

1.1 Needs Analysis and Problem Statement......................................................................... 12

1.2 Project Objectives .......................................................................................................... 13

1.3 Project Approach ............................................................................................................ 14

2.0 Systems Engineering .......................................................................................................... 15

2.1 Team Organization ......................................................................................................... 15

2.2 Needs .............................................................................................................................. 16

2.3 Program Objectives ........................................................................................................ 16

2.4 Schedule ......................................................................................................................... 17

2.5 Project Budget ................................................................................................................ 19

3.0 X-56 Type Design .............................................................................................................. 20

4.0 X-56 Type Fabrication and Assembly ............................................................................... 21

4.1 Fuselage .......................................................................................................................... 21

4.1.1 Material Used .......................................................................................................... 21

4.1.2 Fuselage Fabrication ............................................................................................... 21

4.1.3 X-56 Type Styrofoam Base Repair ......................................................................... 21

4.1.4 Nose Boom with Movable Weight (See Section 5.4) ............................................. 23

4.2 Horizontal Tail ............................................................................................................... 23

4.2.1 Horizontal Tail Fabrication ..................................................................................... 23

4.3 Gimbal and Sting ............................................................................................................ 26

4.3.1 Gimbal Fabrication and Modification ..................................................................... 26

4.3.2 Sting Modification .................................................................................................. 29

5.0 Testing And Preparation .................................................................................................... 30

5.1 X-56 Model Installation On Tunnel Sting With Gimbal Mount (See Reference) ......... 30

5.1.1 X-56 Model Installation Procedure ......................................................................... 30

5.1.2 Model Longitudinal Stability Test Procedures ....................................................... 30

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5.1.2.1 Longitudinal Stability Frequency and Pitch Damping Test ............................ 30

5.1.2.1.1 Pulse Excitation Method Procedure ..................................................................... 30

5.1.2.1.2 Stick-Hit-Nose Boom Excitation Method Procedure ..................................... 31

5.1.3 Test Results (See Section 9.1.2) ............................................................................. 31

5.2 Gust Vane System Installation (See Reference) ............................................................ 32

5.2.1 Test Equipment (See Reference) ............................................................................ 32

5.2.2 Gust Vane Operation Procedure For Varying Vane Frequency and Oscillation

Angle Amplitude ................................................................................................................... 32

5.2.3 Test Plan (See Appendix) ....................................................................................... 32

5.2.4 Test Results (See Section 9) ................................................................................... 32

5.3 Aerodynamic Center Testing and Results ...................................................................... 33

5.3.1 Analysis Procedure ................................................................................................. 33

5.3.2 Test Results ............................................................................................................. 33

5.4 Center of Gravity Testing ............................................................................................... 36

5.4.1 Test Procedures ....................................................................................................... 36

5.4.2 Test Results ............................................................................................................. 38

................................................................................................................................................... 39

5.6 Static Wing Loading Test And Results .......................................................................... 40

5.6.1 Test Procedure ........................................................................................................ 40

5.6.2 Test Results ............................................................................................................. 43

6.0 Theory Predictions Using Athena Vortex Lattice (AVL) .................................................. 45

6.1 Theory Predictions of Models Aerodynamic Center Vs. Center Of Gravity Using

Athena Vortex Lattice (AVL) ................................................................................................... 45

6.3 Theory Predictions Of Model Stability Derivatives Using Athena Vortex Lattice (AVL)

47

7.0 Simulink Real-Time Control System: (See Reference) ..................................................... 49

7.1 Simulink Model Configuration: (See Reference)........................................................... 49

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7.2 The Real-Time Windows Target: (See Reference) ........................................................ 49

7.2.1 Advantages of the Real-Time Windows Target (See Reference) ........................... 49

7.2.2 Known Issues with the Real-Time Windows Target (See Reference) ................... 49

7.3 The X-56 DART Flight Controls Model: (See Reference) ............................................ 49

7.3.1 The ArduPilot Mega(APM) Interface Subsystem (See Reference) ........................ 49

7.3.2 Pilot Input Subsystem (See Reference) ................................................................... 49

7.3.3 The System Status Subsystem (See Reference) ...................................................... 49

7.3.4 The Wind Tunnel Data Recorder Subsystem (See Reference) ............................... 49

8.0 Wind Tunnel Data Acquisition and Analysis .................................................................... 50

8.1 Wind Tunnel Test Data Acquisition (See Reference) .................................................... 50

8.2 Wind Test Data Analysis Method (See Reference) ....................................................... 50

8.2.1 Longitudinal Stability Tests Example Calculations (See Reference) ..................... 50

8.2.2 Gust Response Test Example Calculations............................................................. 50

9.0 Wind Tunnel Test Results .................................................................................................. 51

9.1 Longitudinal Stability Frequency And Pitch Damping Test .......................................... 51

9.1.1 Elevator Pulse Excitation Method Results .............................................................. 51

9.1.2 Stick-Hit-Nose Boom Excitation Method Results .................................................. 51

9.2 Gust Response Magnitude vs. Gust Vane Deflection And Frequency Test- Gust

Response Magnitude vs. Vane Frequency At Various Tunnel Velocities ................................ 52

9.2.1 Test Results ............................................................................................................. 52

9.2.2 Finding Natural Frequency through Gust Response Analysis ................................ 55

9.3 Error and Problems......................................................................................................... 56

10.0 Conclusion ......................................................................................................................... 57

11.0 Recommendation ............................................................................................................... 58

12.0 Contributions and Acknowledgments ................................................................................ 59

13.0 References .......................................................................................................................... 60

13.1 Previous Senior Project Report (Team Falcon by Evan Johnson) ................................. 60

13.2 MATLAT Plots .............................................................................................................. 60

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13.2.1 Stick Excitation Test Plots ...................................................................................... 60

13.2.1.1 Velocity at 77 fps ............................................................................................................. 60

13.2.1.2 Velocity at 80 fps ............................................................................................................. 61

13.2.1.3 Velocity at 90 fps ............................................................................................................. 67

13.2.1.4 Velocity at 100 fps .......................................................................................................... 70

13.2.1.5 Velocity at 110 fps .......................................................................................................... 73

13.2.1.6 Velocity at 120 fps .......................................................................................................... 76

13.2.2 Frequency Excitation Plots ..................................................................................... 79

13.2.2.1 Velocity at 80 fps ............................................................................................................. 79

13.2.2.2 Velocity at 90 fps ............................................................................................................. 81

13.2.2.3 Velocity at 100 fps .......................................................................................................... 83

13.3 Excel Data Results ......................................................................................................... 86

13.3.1 Stick Excitation Test ............................................................................................... 86

13.3.1.1 Velocity at 77 fps ............................................................................................................. 86

13.3.1.2 Velocity at 80 fps ............................................................................................................. 86

13.3.1.3 Velocity at 90 fps ............................................................................................................. 86

13.3.1.4 Velocity at 100 fps .......................................................................................................... 87

13.3.1.5 Velocity at 110 fps .......................................................................................................... 87

13.3.1.6 Velocity at 120 fps .......................................................................................................... 87

13.4 Books .............................................................................................................................. 88

13.5 Other Documents............................................................................................................ 88

13.6 Poster Copy .................................................................................................................... 89

Appendix A-1: Code Listing ........................................................................................................ 90

A-1.1 Matlab Code for short period approximation .................................................................. 90

Appendix A-2: AVL Input Files .................................................................................................. 91

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A-2.1 Geometry File (.avl) ........................................................................................................ 91

A-2.2 Run Case File .................................................................................................................. 95

A-2.3 Airfoil Geometry File: Fuselage.dat ............................................................................... 96

A-2.4 Airfoil Geometry File: Wing.dat ..................................................................................... 99

Appendix A-3: Proposal ............................................................................................................. 102

Appendix A-4: Stick Excitation Test Plan .................................................................................. 110

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List Figures

Figure 4-1 JB Weld KwickWeld................................................................................................... 22

Figure 4-2 Epoxy application on the Styrofoam base. .................................................................. 23

Figure 4-3 Horizontal Stabilizer and Elevator Dimensions .......................................................... 24

Figure 4-4 Vertical Stabilizer Dimensions ................................................................................... 24

Figure 4-5 Horizontal Stabilizer Tail Boom Bracket Mounts ...................................................... 25

Figure 4-6 Modification of the model mount to the gimbal using "bunny ears" graphite epoxy

extensions ...................................................................................................................................... 27

Figure 4-7 "bunny ears" dimension specifications ....................................................................... 27

Figure 4-8 Nose dive due to excessive degree of freedom by the new gimbal configuration ...... 28

Figure 4-9 Nose-down pitch limit bolt on the model's new gimbal configuration ....................... 29

Figure 4-10 Model configuration on the sting mount showing the new carbon fiber extensions 29

Figure 5-1 Excitation Response with Frequency and Damping Calculations .............................. 31

Figure 5-2 AC Planform without tail ............................................................................................ 34

Figure 5-3 AC and CG location vs. tail Longitudinal Location ................................................... 36

Figure 5-4 Test configuration for C.G testing .............................................................................. 37

Figure 5-5C.G. travel with respect to nose weight location for different type of weights ........... 38

Figure 5-6 Top View of the Test Setup ........................................................................................ 40

Figure 5-7 Front View of the Test Setup ...................................................................................... 41

Figure 5-8 Side View of the Test Setup ........................................................................................ 41

Figure 5-9 Tail loaded with 2.5 pounds and with ruler in place to measure deflection ................ 42

Figure 5-10 Set of weights used for experiment ........................................................................... 43

Figure 5-11Graph of Load vs. Deflection for each boom ............................................................. 44

Figure 6-1 Isometric view of the AVL model of the X-56 aircraft .............................................. 45

Figure 6-2 Damping vs. Velocity comparison for theoretical and experimental .......................... 48

Figure 9-1 Pitch Rate Damping versus Velocity .......................................................................... 51

Figure 9-2 Pitch Rate Frequency versus Velocity ........................................................................ 52

Figure 9-3 Pitch Rate Damping Response with Gust Vane oscillating at 1.25Hz ........................ 53

Figure 9-4 Pitch Rate Damping Response with Gust Vane oscillating at 1.5Hz .......................... 53


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Figure 9-11 Model Maximum Peak to Peak Oscillating Pitch Amplitude vs. Gust Vane

Frequency @ V = 90 ft/s ............................................................................................................... 56

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List of Table

Table 2-1Project Expenditures ...................................................................................................... 19

Table 5-1 The values inputted into the VFD and its corresponding characteristics on the

flywheel......................................................................................................................................... 32

Table 5-2 Weight breakdown and calculations of new C.G with respect to nose weights location

....................................................................................................................................................... 39

Table 5-3 Load vs Horizontal Tail Trailing Edge Deflection Data (left and right labeled as seen

with aircraft upside-down) ............................................................................................................ 44

Table 6-1 Predicted stability derivatives with respect to the horizontal tails distance ................. 46

Table 6-2 Variation of cmα for different velocities using AVL ................................................... 47

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1.0 INTRODUCTION

1.1 NEEDS ANALYSIS AND PROBLEM STATEMENT 1.1.1 Next generation aircrafts are implementing higher AR wings with high flexibility

(Blended Wing Bodies, UAV, Strut Braced Wings, etc.). These wings are prone to

experiencing high dynamic deflection and high stress for gust response and flutter that

could lead to structural failure and unacceptable flying qualities. The NASA/ Lockheed

Martin X-56 is an example of such aircraft. This project scales by design failures of the X-

56 into a ‘free-flying” wind tunnel model. Further research on aero-elastic phenomena,

specifically the analysis of flutter and gust response challenges, will contribute to creating

successful next generation aircrafts.

1.1.2 Of the four major aerospace engineering disciplines, control systems offers the least

undergraduate courses despite having become one of the most important disciplines in

industry. This project allows for a greater understanding of advanced control topics not

covered in any undergraduate course.

1.1.3 If the department were to include practices and test procedures from this project in the

aerospace curriculum in the form of a lab, undergraduate students would be exposed to

data acquisition and controls hardware.

1.1.4 Students will be provided with hands on experience with testing a blended wing-body’s

Stability Augmentation System (SAS).

1.1.5 Explore and test alternative control methods for a blended wing-body type UAV aircraft.

Utilize differential rudder and aileron actuation in order to aid in yawing motion through

differential drag effects.

1.1.6 Demonstrate the feasibility of a remotely processed control system running on a portable

computer running a real-time Simulink control system.

1.1.7 Provide the Aerospace Engineering department with a fully functional flutter research

model capable of modeling advanced control concepts aimed at actively suppressing aero-

elastic flutter during simulated flight in the low-speed wind tunnel.

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1.2 PROJECT OBJECTIVES 1.2.1 Modifying the blended wing body model fabricated by the 2013-2014 FALCON Club

senior project team. The modified scale model of the X-56 will have a relocated C.G by

the use of a nose boom weight in front of the model. The goal of this is to move the AC aft

of the C.G. for static and dynamic stability.

1.2.2 The FALCON (Flutter ALleviation and CONtrol) model will be tested in the subsonic

wind tunnel with two degrees of freedom for longitudinal stability testing in order to

develop an optimized longitudinal Stability Augmentation control system.

1.2.3 Demonstrate a longitudinal gust alleviation system capable of reacting to vertical gusts in

the subsonic wind tunnel. This test will utilize the gust generation system designed and

installed in the wind tunnel by the 2012-2013 Flutter Club team.

1.2.4 Expand stability augmentation system to include lateral-directional motion. Demonstrate

controllability and augmented static and dynamic stability of a blended wing-body aircraft

in five degree of freedom motion.

1.2.5 Super impose the stability augmentation system with gust alleviation system to reduce

flutter phenomena for a rigid wing structure.

1.2.6 Perform preliminary vibration and flutter analysis in NASTRAN for a composite-skinned

flexible wing optimized for span wise torsional bending. Determine optimal material,

wing structure, and mass distribution to obtain desired structural dynamic modes.

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1.3 PROJECT APPROACH 1.3.1 Approximate X-56 type model CG and AC locations. Design tail that will move AC

further back while not affecting CG location. Use light material for tail

1.3.2 Create Horizontal Tail with RC controlled elevator for enhanced trim control and

longitudinal stability to add to Existing X-56 Model

1.3.3 Add functioning elevator to control pitch

1.3.4 Lock wingtip rudders

1.3.5 Use Cal Poly Low Speed Wind Tunnel to achieve stable flight

1.3.6 Attach model to gimbal mount, then attach model/gimbal mount apparatus to crescent

sting in order to simulate free flight in pitch and plunge.

1.3.7 Use Team Falcon’s Simulink model in order to control model with joystick

1.3.8 Use the on board gyros and accelerometer to measure pitch, pitch rate, and time

1.3.9 Simulate turbulent wind conditions by fluctuating the model in pitch. This can be

accomplished by performing an elevator pulse-doublet as well as by tapping the nose

boom with a rod.

1.3.10 Vary gust vanes frequency in the tunnel in order to find the model’s natural short period

frequency and maximum response amplitude for future use in designing a gust alleviation

system

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2.0 SYSTEMS ENGINEERING

2.1 TEAM ORGANIZATION

ProgramManagerTuanDinhJr

(909)[email protected]

LeadControlSystemsEngineer

GeorgePaguio(213)304-5036

[email protected]

ControlSystemsEngineer

DongJinRyoo(714)873-2572

jus [email protected]

LeadAerodynamicsEngineerBillWogahn


AerodynamicsEngineer

ReginaldGuinto(909)539-5090

[email protected]

AerodynamicsSupport

CrystalNunez(714)619-1434

[email protected]

AerodynamicsSupport

AnahiHernandez(323)490-2959

[email protected]

LeadStructuralEngineerTannerClark


StructuralEngineerJasonKong

ChiefFabrica onEngineerJasonKong


Fabrica onEngineerTannerClark

ChiefFinancialOffic

e

r

AryaWilliams(626)710-0219

[email protected]

DeputyProgramManagerDwightNava


ProgramandControlSystems

AdvisorEvanJohnson(714)851-4146

[email protected]

FacultyAdvisorStevenDobbs

CalPolyPomona2014-2015SeniorProject–TeamGUAC

GustAlleviationoftheDartX-56

Figure 2.1-1 Team GUAC Organization Chart

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2.2 NEEDS Research towards flutter alleviation of blended wing bodies is few and far between. The

study of gust alleviation and stability augmentation systems of blended wing bodies is to this day

one of the subjects being actively studied in industry today.

For the team, this projects provides team members with the opportunity to be exposed to the

systems engineering process. By integrating aerodynamic theory to application and applying the

manufacturing process, students are introduced through the design life cycle and manage the

entire program with the principles of system engineering.

2.3 PROGRAM OBJECTIVES This is a multi-year research opportunity exploring the control alleviation of gusts of

blended wing bodies. For the fourth year iteration of the project, the goal of the GUAC Team is

to make the aircraft Aerodynamically Stable. Previous year’s design had the center of gravity aft

the aerodynamic center which contributed to its stability issues. To correct these issues GUAC

will modify the design by incorporating a twin-boom tail, including a new elevator and two new

rudders to the aft section of the model. Thus correcting model’s stability issues.

By developing a stability as well as a gust alleviation system for a rigid wing model, this

provides a baseline to explore other autopilot and stability augmentation systems of flexible

systems for future teams to research. The mechanics to control the magnitude of the gust had

been set in place. However the capabilities of the gust-vane system had not been measured and

thus after the previous year had just finished constructing the gust-vane system and to pick up

from last year, the gust-vane system flow dynamics must be characterized.

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2.4 SCHEDULE

Figure 2.4-1 Team GUAC Schedule 1

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Figure 2.4-2 Team GUAC Schedule 2

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2.5 PROJECT BUDGET Since the team had no initial funding, the team had to provide their own funding and take alternative

measures to reduce cost. Beginning of 2014 Fall Quarter the team applied for the Kellogg Future Mini-

Grant 2014-2015 Program for the amount of $1,400, but the team was denied later that quarter. With

help from Cal Poly Pomona’s solar boat team, donated carbon fiber plate was used in the fabrication

process of the gimbal and sting modifications. In addition California Space Grant Consortium, they have

donated $330 to the team. Located below in Table 2.5-1, displays the 2014-2015 expenditures.

Table 2.5-1 Project Expenditures

Item Quantity Category Cost Per Item Total Cost

2 Sheets of Baltic Birch 3mmx 2 Raw Materials $4.45 $9.70

Carbon Fiber Rod 4 Raw Materials $7.99 $48.03

Mach Screws and Washers 4 Materials $1.18 $4.72

Aluminum Sheet 1 Raw Materials $24.08 $26.02

Square Carbon Tube 2 Raw Materials $8.39 $22.27

Servo Extension Cable 2 Electronics $5.99 $13.18

Thumb Drive 1 Electronics $12.99 $14.16

Binder 1 Materials $3.99 $4.50

Airtronics 94802 Sub-Micro Digital BB Servo 2 Electronics $35.99 $71.98

Rosin Core Solder Minit 1 Electrical Equipment $2.99 $4.19

J-B Kwikweld 3 Materials $5.27 $16.91

Parallels For Mac 1 Software $20.00 $20.00

Dremel Bit Sizes 2 Hardware $3.99 $9.33

Piano Hinges 2 Materials $1.99 $4.30

Nylon Hinges 1 Materials $6.70 $6.70

PosterBoard 1 Materials $15.79 $17.21

Total: $293.20

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3.0 X-56 TYPE DESIGN After the center of gravity and the aerodynamic center of the X-56 were located, it was discovered

that the reason the model was unstable in previous wind tunnel tests was because of the location of the

aerodynamic center with respect to the center of gravity. The best way to correct this problem and move

the aerodynamic center was to install a horizontal tail using two booms, which would be installed to the

body of the aircraft. It was determined that the body of the model has to be modified in order to make

room for the two booms as well as the assembly to hold the booms. This modification is done by carving

out some of the foam in the shape of the aluminum “U” brackets on the interior of the bottom half of the

body using a Dremel Rotary Tool. The back of the body will also have to be carved out in order for the

booms to protrude from the back of the aircraft. After the body is shaped to attach the tail booms, all

rough edges will be sanded down in order to create a smooth surface for the air to flow over during the

wind tunnel testing.

While these modifications are being done to the fuselage, the CAD model of the fuselage will be

modified in order to create a new fuselage, which would allow for the tail booms. After the CAD model

is modified, a 3-D printer will be used to create a negative of the base. This negative will then be used to

create a mold of the body of the aircraft. Using this mold, a solid base can then be created using the

same material as the original body for consistency. After the mold is completed, the exterior will be

sanded down to create as smooth of a surface as possible for the air to flow over it. However, fabrication

of this new fuselage was not performed in this project.

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4.0 X-56 TYPE FABRICATION AND ASSEMBLY

4.1 FUSELAGE For an aircraft to be aerodynamically stable, its aerodynamic center should be located behind the

center of gravity. However, in the X-56’s current design and configuration, the center of gravity is well

aft of the aerodynamic center by about 2.25”. In order for the aircraft to become aerodynamically stable,

the proposal is to incorporate a twin-boom tail, including a new elevator and two new rudders, to the aft

section of the aircraft. This in turn should help shift the aerodynamic center of the aircraft aft and help

with aircraft stability.

4.1.1 Material Used

2 Carbon fiber rods, minimum of 18” in length and 3/8” in diameter

1 medium-sized birch plywood board, 1/16” thick

2 Airtronics 94802 Sub-Micro Digital BB Servos

Actuators (2 rudders and 1 elevator)

4.1.2 Fuselage Fabrication Two carbon-fiber square rods have been purchased from Rockwest Composites, each stick

measuring 2 feet in length and 3/8th inches in diameter. A piece of birch plywood measuring 1/16” thick

was also purchased for crafting the new vertical and horizontal stabilizers and twin vertical tails. The

two booms will be internally mounted to the aircraft’s carbon fiber skeleton via four aluminum “u”

bracket mounts, two on either side of the aircraft. The booms will be able to slide along its mount to a

specified length based on performance, and can be held in place with a lock screw. Since there are new

protrusions coming out of the aircraft’s fuselage, the current Styrofoam base of the fuselage either needs

to be remade or modified to fit the new tail booms while maintaining the aircraft’s aerodynamics.

4.1.3 X-56 Type Styrofoam Base Repair After the team finished the first wind tunnel test, the Styrofoam base of the model saw cracks in

three locations causing the entire base to break into multiple pieces. All the cracks were due to the high

amount of stress directed at the Styrofoam base whenever the aircraft experienced rough movement

while the team was trying to get used to the sensitivity of the aircraft’s control surfaces in the tunnel.

Because most glues tend to decompose or corrode Styrofoam, the choice was made to use quick-setting

steel reinforced epoxy. The choice of purchase was JB Weld’s KwikWeld, and is shown in Figure 4.1.3-

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1 below. The epoxy sets in about four minutes and cures in approximately 4 hours; with a listed shear

strength of 2424 psi.

Figure 4-1 JB Weld KwickWeld

To begin repairs, the Styrofoam base was removed from the model. Some of the Styrofoam at the

area of impact was missing, presumably having been either scattered on the ground or blown down the

length of the wind tunnel. Therefore, the epoxy was also used to fill in the gaps. To fix the rear section

of the Styrofoam base that is right underneath the aircraft’s body flap, a piece of duct tape was used to

initially hold the piece together. A piece of cut carbon fiber as well as double layer of epoxy was applied

across the length of the section. This, as well as the epoxy used to fill the side of the base where the

Styrofoam was carved out to house the gimbal mount, can be seen in Figure 4.1.3-2.

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Figure 4-2 Epoxy application on the Styrofoam base.

To ensure that the double layer of epoxy was fully set, the Styrofoam base was left untouched for 24

hours, and was then reattached to the model the following day. Tests were subsequently performed for

the rest of the week without incident, and no further stress cracks were noticed on the repaired base.

4.1.4 Nose Boom with Movable Weight (See Section 5.4)

4.2 HORIZONTAL TAIL

4.2.1 Horizontal Tail Fabrication Construction of the tail began by acquiring a piece of birch plywood that was 1/8” thick. The

piece was then cut into four pieces; two vertical stabilizers, a horizontal stabilizer and an elevator whose

dimensions can be seen below in Figures 4.2.1-1 and 4.2.1-2. After the pieces were all cut, the edges

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were sanded using course sandpaper to give them the shape of an airfoil, and then using fine sand paper

to smooth the surfaces and decrease drag.

Figure 4-3 Horizontal Stabilizer and Elevator Dimensions

Figure 4-4 Vertical Stabilizer Dimensions

8”

7.25”

2”

1.5”

1.125”

1.875”

3.375”

25

The assembled tail was then coated with a wood primer and painted. After the paint dried, the

pieces were sanded again and repainted to make the surfaces as smooth as possible. While the tail was

drying for the final time, the tail booms were constructed. A piece of ¼” carbon square tubing was cut

into two 24” long pieces. From there, the two square carbon tubes were measured out so that the new tail

would not collide with the wind tunnel sting mount, but long enough to assist in shifting the

aerodynamic center back as well as providing a bigger moment arm for the control surfaces on the new

tail. To mount the booms, aluminum brackets were molded from strips of sheet aluminum. Holes were

drilled into the carbon-fiber skeleton of the model, and the aluminum brackets were mounted onto the

body. The carbon tubes were then inserted into the brackets and tightened by clamping the brackets with

two bolts on each bracket, one on each side. Two brackets were used for each square carbon tube.

To mount the new tail onto the square carbon tubes, two separate aluminum brackets were made.

They were then glued onto the bottom of the horizontal stabilizer with epoxy with the carbon tubes,

shown in Figure 4.2.1-3 below.

Figure 4-5 Horizontal Stabilizer Tail Boom Bracket Mounts

Any excess carbon tubing was removed to reduce weight and drag on the aircraft. Finally the

elevator was mounted onto the horizontal stabilizer with four hinges spread out across the length of the

control surface.

To ensure a more snug fit, two grooves were created in the Styrofoam base of the aircraft to

allow the two new tail booms to slide into the base without leaving too much of a gap between the upper

26

and lower parts of the fuselage. Finally, a new servo and servo arm was attached to the bottom surface

of the horizontal stabilizer to connect to the new elevator. A small rounded foam mold was made to fit in

front of the servo to minimize drag. The original wire connecting the servo to the Ardupilot was not long

enough, so a 24-inch extension cable was purchased and used to complete the connection.

4.3 GIMBAL AND STING

4.3.1 Gimbal Fabrication and Modification After last year’s recommendation and visual inspection by our advisor this year, we decided that

the gimbal assembly needed to be modified. The gimbal was initially designed for the 3D model of team

prior to the senior project team last year (FALCON). FALCON reused the gimbal was due to time

constraint and machining experience. For this year, no one had machining experience nor found anyone

who could machine a new gimbal, therefore it was decided to just modify the current gimbal.

The first modification made was to find a way to properly mount the gimbal to the frame of the

model. Since the original design accounted for a variable location for the gimbal and we were not able to

reconfigure the model, there was an issue with gap being too far apart for the screws to fit both sides.

This was resolved by using a 4 inch bolt that goes the frame and the gimbal with collars and spacers in

between the two sides. This configuration can be seen in Figure 4.3.1-1. It was secured by putting in two

nuts at the end of the bolt because the vibration would unscrew the nut if it were just one bolt.

After further inspection, Professor Dobbs noticed that the translation degree of freedom in the

negative direction was limited because the gimbal would simply hit the base plate. We resolved this by

extending the mounts using ‘bunny ears’. The bunny ears configuration is shown in Figure 4.3.1-1. The

ears were placed as close to the center of gravity as much as possible but due to the pre cuts, we were

able to just mount the ears as far back as possible on the slots.

27

Figure 4-6 Modification of the model mount to the gimbal using "bunny ears" graphite epoxy extensions

The bunny ears are made out of a 13 layer carbon fiber flat plate. The Dimensions are shown on

Figure 4.3.11-2:

Figure 4-7 "bunny ears" dimension specifications

As we tested in the wind tunnel we found flaws in our gimbal modification designs. As soon as

the pivot point reaches its extended peak (the sting mount collinear with the “bunny ears”, the model

would pivot towards the front and the model would noise dive as shown on Figure 4.3.1-3.

Slot for variable

mounting

Gimbal

Sting

mounting

Bunny

Ears

28

Figure 4-8 Nose dive due to excessive degree of freedom by the new gimbal configuration

This was a serious risk especially when trying to run the tunnel at higher speeds. The nose dive

occurred twice and broke the fuselage foam in half. We resolved this risk by adding additional bunny

ears that held a secondary bolt that goes in between the gimbal. This configuration is shown in Figure

4.3.1-4. The purpose of the bolt was to act as a stopper and to limit the translation travel of the gimbal.

The translation is now limited to when the bolt hits the top and bottom part of the gimbal. This design is

ideal because this would notify us when the plane is at trim. As the wind pull the model in the direction

of drag the stopper bolt would hit against the top and bottom part of the gimbal due to change in lift and

when the plane is in trim, the bolt is neither hitting the top or bottom.

29

Figure 4-9 Nose-down pitch limit bolt on the model's new gimbal configuration

4.3.2 Sting Modification The sting was modified last year to elevate the model since the sting was also initially designed

for the previous’ model. FALCON used aluminum blocks as their extended leg. We thought that these

blocks where too big and non-aerodynamic so therefore we recreated them using a 13 layer carbon fiber

plates. This new sting mount is shown in Figure 4.3.2-1

Figure 4-10 Model configuration on the sting mount showing the new carbon fiber extensions

New carbon fiber

extensions

Added bolt to limit the

degree of freedom in

the vertical motion

30

5.0 TESTING AND PREPARATION

5.1 X-56 MODEL INSTALLATION ON TUNNEL STING WITH GIMBAL

MOUNT (SEE REFERENCE)

5.1.1 X-56 Model Installation Procedure Installation of the X-56 model with added horizontal tail was very similar to previous year’s

methods of installation. After modifying the gimbal (see gimbal modification section), the threat of

nose-diving during testing was addressed, and the model could be installed onto the sting of the wind

tunnel. Before mounting the gimbal onto the sting, thin strips of hinge tape were applied to the end of

the sting that meets the gimbal. This helped to secure the gimbal and also protect the material of the

sting. Once the gimbal was in place, two metal hose clamps were used to secure the gimbal firmly onto

the sting. The model was then installed and ready for its wiring to be secured. The wiring from the

model to both the power source and the computer were ran down the gimbal and onto the sting in order

to exit the wind tunnel. The wiring was secured by electrical tape and was taped down as securely as

possible. Both a front view and a rear view of the model can be seen installed in the wind tunnel in

figures XX and XX. After exiting the wind tunnel, the power wire ran to the power source provided by

Cal Poly Pomona Engineering Department. The cable from the Arduino board was connected to the

computer used for data analysis and Simulink control. The model was then ready for testing. For a more

in depth report of model installation, see previous years report.

5.1.2 Model Longitudinal Stability Test Procedures Wind Tunnel testing for the D.A.R.T. model this year was centered on gathering pitch rate data

of the model at different velocities in order to determine a pitch rate frequency damping coefficient. To

accomplish this, the controls module ran through MATLAB’s Simulink was modified to activate only

the new horizontal tail that was fabricated this year. This meant that pitch was the only controllable

movement of the model when testing in CPP’s subsonic wind tunnel.

5.1.2.1 Longitudinal Stability Frequency and Pitch Damping Test

5.1.2.1.1 Pulse Excitation Method Procedure

The Elevator Pitch Excitation is obtained while able to maintain trim flight. After maintaining

trimmed flight, the pilot, then excite the model to simulate a pitch doublet.

31

5.1.2.1.2 Stick-Hit-Nose Boom Excitation Method Procedure

During testing, the model was trimmed to simulate flight at different velocities. Once set at

stable trim, a pitch pulse was simulated by tapping the nose of the model with a thin and slender rod

through the top slit in the wind tunnel ceiling. This pitch pulse was simulated 6 times at each velocity,

and pitch rate versus time data was gathered through the Arduino board set in the model.

The damping coefficient was determined through Equation 5.1.2.1.2-1:

𝑔 = 1

𝑛𝜋𝐿𝑛(

𝐴𝑜

𝐴𝑛)

Equation 5.1.2.1.2-1

Where n is the number of amplitude cycles within the simulated pitch pulse and 𝐴𝑜 is the

magnitude of the initial amplitude and 𝐴𝑛 in the cycle amplitude at nth cycle this equation was applied

for each of the 6 simulated pitch pulses per velocity. The frequency was calculated by counting the

number of cycles over a time period and dividing into cycles per second as shown Figure 5.1.2.1.2-1.

Figure 5-1 Excitation Response with Frequency and Damping Calculations

To accurately determine the logarithmic damping value, the “Ln A” was plotted vs. ‘n’ and a

straight line curve fit was made to use in determining 𝐴𝑜 and 𝐴𝑛.

5.1.3 Test Results (See Section 9.1.2)

ΔT = 2.9s

2 1

An = .45

A0 = 1.0

3 𝐹𝑟𝑒𝑞𝑢𝑒𝑛𝑐𝑦 = 3 𝐶𝑦𝑐𝑙𝑒𝑠

2.9 𝑆𝑒𝑐 = 1.03 Hz

Damping: 𝑔 = 1

3𝜋𝐿𝑛(

1.0

.45) = .08447

0

32

5.2 GUST VANE SYSTEM INSTALLATION (SEE REFERENCE)

5.2.1 Test Equipment (See Reference)

5.2.2 Gust Vane Operation Procedure For Varying Vane Frequency and Oscillation

Angle Amplitude To operate the Gust Vane, One must first have installed and plugged the motor that would later

translate toward the Gust Vanes within the Tunnels. The frequency is then manually changed through

the use of the motors control on top of the Low Speed Wind Tunnel. For the set-up of the frequency, the

table below will showcase the relations between the rotations per second of the motor to frequency.

Table 5-1 The values inputted into the VFD and its corresponding characteristics on the flywheel.

VFD RPM Flywheel RPM Flywheel RPS/Frequency (Hz)

500 50 0.83

750 75 1.25

900 90 1.5

1050 105 1.75

1200 120 2

1500 150 2.5

1500 150 3

2100 210 3.5

2400 240 4

2700 270 4.5

3000 300 5

5.2.3 Test Plan (See Appendix)

5.2.4 Test Results (See Section 9)

33

5.3 AERODYNAMIC CENTER TESTING AND RESULTS

5.3.1 Analysis Procedure The aerodynamic center is the point where pitching moment coefficient does not vary with lift

coefficient. This makes it the point where the lift acts on an airfoil or where the total lift acts on a whole

aircraft. Because of this, the total moment about the nose of the plane can be represented by

𝛴𝑀𝑜 = 𝐿𝑡𝑜𝑡 ⋅ 𝐴𝐶𝑡𝑜𝑡

where Ltot is the total lift of the plane and ACtot is the plane’s aerodynamic center measured from

the nose. The total moment can also be described as the sum of the products of lift and AC location for

each individual component of the plane, thus making the moment about the nose reference point:

𝛴𝑀𝑜 = 𝐿𝑡𝑜𝑡 ⋅ 𝐴𝐶𝑡𝑜𝑡 = ∑ [𝐿𝑖 ⋅ 𝐴𝐶𝑖]𝑛𝑖=1 ,

with n being the total number of components that change lift due to an angle of attack change.

From here, the lift variable can be represented by its definition,

𝐿 = 𝐶𝑙⍺ ⋅ ⍺ ⋅ 𝑞 ⋅ 𝐴,

where Cl⍺ is the coefficient of lift to angle of attack, ⍺ is the angle of attack, q is the dynamic

pressure, and A is the planform area. Once this is added to the original equation, it reads:

𝛴𝑀𝑜 = (𝐶𝑙⍺𝑡 ⋅ ⍺ ⋅ 𝑞 ⋅ 𝐴𝑡𝑜𝑡) ⋅ 𝐴𝐶𝑡𝑜𝑡 = ∑ [(𝐶𝑙⍺𝑖 ⋅ ⍺ ⋅ 𝑞 ⋅ 𝐴𝑖) ⋅ 𝐴𝐶𝑖]𝑛𝑖=1 .

Since angle of attack and dynamic pressure are assumed to be independent of component

location, they can be removed from the summation and canceled out through division on both sides of

the equation. By moving Cl⍺t and Atot to the other side of the equation, the final result for the calculation

of the aerodynamic center location can be shown as

𝐴𝐶𝑡𝑜𝑡 =∑ [(𝐶𝑙⍺𝑖 ⋅ 𝐴𝑖) ⋅ 𝐴𝐶𝑖]𝑛

𝑖=1

𝐶𝑙⍺𝑡 ⋅ 𝐴𝑡𝑜𝑡

5.3.2 Test Results After having solved for the aerodynamic center equation for the entire aircraft, the variables A,

Cl⍺, and AC for the individual components needed to be found. For simplicity, the plane was divided in

half along the center and then broken into 3 separate components: the fuselage, the inboard wing

segment, and the outboard wing segment, as shown in Figure 5.3.2-1 below.

34

Figure 5-2 AC Planform without tail

For each of the segments’ areas, an approximated drawing of the plane’s planform (similar to the

one shown in Figure 5.3.2-1) was created in order to determine the areas graphically. Along with the

segment areas, the aerodynamic center locations could also be determined with a planform drawing by

finding the approximate Mean Aerodynamic Chord (MAC) location along the span of each segment and

then measuring the quarter chord distance along the MAC. This was accomplished by first measuring

the length of the tip and root chords on the plane and then drawing them to scale. The length of the root

chord was added once to each end of the tip chord and vise-versa. By connecting the diagonals of the

newly extended ends, the location of the mean aerodynamic chord (MAC) could be found by

pinpointing the intersection between the two diagonals. Since the AC is assumed to be located at a

quarter of the MAC length for subsonic aircraft, the aerodynamic centers for the first two segments were

able to be found by adding the vertical distance from the nose to the wing tip at the MAC to the quarter

chord of the MAC. For the outboard wing, it was found that the MAC was located at the midpoint of the

segment (in the spanwise direction) since the root and tip chords were equal to each other. By adding the

vertical distance from the nose, the AC for the third segment was found.

35

Once the individual areas and AC locations were found, the Cl⍺ for each component was solved

for. For most straight wings and airfoils, the Cl⍺ can be approximated as 2𝜋, however for the swept

segments of the plane, the Cl⍺ had to be found by multiplying the cosine of the sweep angle by 2𝜋.This

applied to the inboard and outboard wing segments, while the fuselage was merely approximated at 2𝜋.

Lastly, the Total area and the total aircraft Cl⍺t had to be found. The total area was determined by

adding up the individual areas and then multiplying by two since the model was split in half. The Cl⍺t

could not be calculated graphically or by assuming a value of 2𝜋. It needed to be found through testing

and analysis. For this, the aerodynamic performance program AVL was used in order to find the value

for Cl⍺t. Given all of the values, the AC could then be found by plugging everything in and multiplying

by 2 in order to account for both halves of the plane.

When the tail was considered, a fourth component had to be added to the aerodynamic center

equation that took into account the tail distance, area, and shape. First, a standard rectangular shaped tail

was selected that had a span of about 5.53 inches and a chord length of 4 inches. Being rectangular and

centered about the central axis of the plane, the Cl⍺ was assumed to be 2𝜋, the MAC was determined to

be directly in the middle of the span, and the aerodynamic center was assumed to be at the ¼ chord. The

actual AC distance from the nose of the plane was not yet decided, so several distances were tested,

ranging from 2 inches from the trailing edge of the plane to well over a foot away. Results for AC

location vs Tail were compiled and graphed along with current CG vs tail data to show tail distances that

would produce stable results, as shown in Figure 5.3.2-2 below.

36

Figure 5-3 AC and CG location vs. tail Longitudinal Location

5.4 CENTER OF GRAVITY TESTING Knowing that the model experienced pitch unstable flight(static divergence) in the wind tunnel in

last year’s wind tunnel tests, the location of the center of gravity needed to be determined. This test for

the center of gravity allowed for a rough estimate for the C.G. location.

5.4.1 Test Procedures Due to the lack of equipment at the time, a makeshift C.G. locator was created using a plastic

soda bottle with curved cap. The curved top of the cap allowed the model to be balanced on a semi-fine

point.

To mark the location of the center of gravity on the model, hash marks were drawn onto the

fuselage of the model with each hash mark representing 1/4th of an inch. The zero marker was recorded

with a piece of tape on the nose.

The initial center of gravity without the attachment of the nose weight came out to be 9.5 inches

0

2

4

6

8

10

12

0 10 20 30 40

Lo

cati

on

Fro

m N

ose

(in

)

Tail Distance (in)

AC Location and CG Location vs. Tail Distance

AC Location

CG Location

37

behind the zero marker on the nose. There were 4 available nose weights to install into the nose boom

and Figure 5.4.1-1 shows the configuration for this test. Due to the risk of making the airplane too heavy

to fly, we chose the 190 g weight for testing.

Figure 5-4 Test configuration for C.G testing

XNose Weight

Reference point

0

XNew C.G.

Center of gravity

without nose weight

38

5.4.2 Test Results

Figure 5-5C.G. travel with respect to nose weight location for different type of weights

Figure 5.4.2-2 shows the different types of weight used and its corresponding effect on the C.G.

of the model. We decided to choose the 190g and place it 12 in in front of the nose. The new C.G based

on the calculations was found to be 7.95 in from the nose where the actual C.G. came out to be 8.125 in

from the nose when measured by hand. Table 5.4.2-1 displays a complete breakdown of weights and

distances of the nose weights for this test.

0

2

4

6

8

10

12

0 1 2 3 4 5 6 7 8 9 10 11 12 13 14

Dis

tan

ce o

f w

eigh

t fr

om

th

e n

ose

(in

)

C.G of the aircraft (in)

CG Travel with Respect to Nose Weight Location

228 g

190 g

146 g

120 g

W/out noseboom

39

Xcg of plane 9.5 in Wplane 3.584712 lbs Wweight 228 g 0.50265336 lbs Wtotal 4.356329 lbs

Lfuselage 14.796 in WH.T 0.268964 lbs 190 g 0.4188778 lbs 4.272554 lbs

Xdistance of H.T. 3.5 in 146 g 0.32187452 lbs 4.17555 lbs

Xcg actual 8.125 in 120 g 0.2645544 lbs 4.11823 lbs

Current

Xdistance of weight in boom (in) Xcg (in) Xdistance of weight in boom (in) Xcg (in) Xdistance of weight in boom (in) Xcg (in) Xdistance of weight in boom (in) Xcg (in) x x

1 8.831534 1 9.02431 1 9.257187 1 9.399953 1 9.5

2 8.71615 2 8.92627 2 9.180101 2 9.335713 2 9.5

3 8.600765 3 8.828231 3 9.103016 3 9.271473 3 9.5

4 8.485381 4 8.730192 4 9.02593 4 9.207233 4 9.5

5 8.369996 5 8.632153 5 8.948845 5 9.142994 5 9.5

6 8.254611 6 8.534114 6 8.871759 6 9.078754 6 9.5

7 8.139227 7 8.436074 7 8.794674 7 9.014514 7 9.5

8 8.023842 8 8.338035 8 8.717588 8 8.950274 8 9.5

9 7.908457 9 8.239996 9 8.640503 9 8.886034 9 9.5

10 7.793073 10 8.141957 10 8.563417 10 8.821794 10 9.5

11 7.677688 11 8.043917 11 8.486332 11 8.757555 11 9.5

12 7.562304 12 7.945878 12 8.409246 12 8.693315 12 9.5

228 190 146 120

Table 5-2 Weight breakdown and calculations of new C.G with respect to nose weights location

40

5.6 STATIC WING LOADING TEST AND RESULTS

5.6.1 Test Procedure 5.5.1.1 The purpose of the horizontal tail loading test was to assure that that the tail would not

structurally fail during the wind tunnel tests. The predicted maximum load for the tail was 9.5

lbs. This was determined by assuming a CLαtail of 2π, a maximum angle of attack of 14 degrees

and a maximum test velocity of 120 feet per second. Then Ltailmax = CLααmax(0.5ρVmax2)*Atail,

where the area of the tail is about 27.5 square inches. The foam fuselage and top cover was

removed. The plane is laid upside down and supported on two aluminum blocks. Two clamps

are placed on the front part of the carbon fiber skeleton of the aircraft. The test setup is shown

in the Figures 5.5.1-1, 5.5.1-2, and 5.5.1-3.

Figure 5-6 Top View of the Test Setup

41

Figure 5-7 Front View of the Test Setup

Figure 5-8 Side View of the Test Setup

5.5.1.2 The deflection of the tail was measured using a 1 foot ruler with 1/32-inch accuracy.

5.5.1.3 Loading method

5.5.1.3.1 Location

Between and not touching booms

42

Enable tail attachment and boom attachment strength

Figure 5-9 Tail loaded with 2.5 pounds and with ruler in place to measure deflection

43

5.5.1.3.2 Weight Loading

The weights incremented by 3/4s of a pound.

Half the weight on each side in order to distribute the load

Figure 5-10 Set of weights used for experiment

5.5.1.4 Data & Results

Plot boom tip deflection vs. load

Look for non-linear slope, indicating failure

5.6.2 Test Results The test results are given in Table 5.5.1-1 and Figure 5.5.2-1

44

Table 5-3 Load vs Horizontal Tail Trailing Edge Deflection Data (left and right labeled as seen with aircraft upside-down)

Figure 5-11Graph of Load vs. Deflection for each boom

As seen in Figure 5.5.2-1, the Horizontal Tail Deflection vs Load approximated a straight line all

the way to the maximum load of 1.75*2 = 3.5 lbs. This means that there was no structural damage

occurred due to the fact that the slope of the line would have decreased if a crack occurred due to the

reduced stiffness. Therefore, it was determined that the tail would remain structurally sound up to the

planned maximum testing conditions for the wind tunnel tests.

load per side (lb) load (lb)

left boom right boom left boom right boom

0 0 3.4375 3.625 0 0

0.5 1 3.40625 3.5625 0.03125 0.0625

1 2 3.34375 3.59375 0.09375 0.03125

1.75 3.5 3.3125 3.5 0.125 0.125

2.5 5 3.25 3.375 0.1875 0.25

3.25 6.5 3.1875 3.34375 0.25 0.28125

4 8 3.15625 3.21875 0.28125 0.40625

4.75 9.5 3.09375 3.1875 0.34375 0.4375

deflection (in)

45

6.0 THEORY PREDICTIONS USING ATHENA VORTEX

LATTICE (AVL)

6.1 THEORY PREDICTIONS OF MODELS AERODYNAMIC CENTER VS.

CENTER OF GRAVITY USING ATHENA VORTEX LATTICE (AVL) AVL was used to generate theoretical values for the stability derivatives of the FALCON model

with and without the horizontal tail. These longitudinal and lateral-directional stability derivatives can

used to design a stability augmentation system through the use of state space modeling in future work.

This year, modifications to the geometry and mass property codes were implemented to achieve more

accurate values of stability derivatives. The AVL model of the X-56 type aircraft is show in Figure

5.6.2-1 shown below.

Figure 6-1 Isometric view of the AVL model of the X-56 aircraft

With a full 3-D aircraft model in solid works, the mass properties (i.e. Moment of inertias,

46

weight, and center of gravity) were obtained and are implemented as a mass file in AVL. A center of

gravity test was also conducted to measure the C.G experimentally. The experimental value was

measured to be 8.125 in aft of the nose with the horizontal tail installed and 190 gram on the nose boom

12 in forward of the nose. In contrast, the solid works mass property feature measures the C.G of the

model to be at 9.29 in. The experimental value was used in the AVL model due to the fact that

experimental results yield higher accuracy over tools such as solid works because it does not account for

every detail of the model.

Due to the longitudinal instability of the model and AVL model, a horizontal tail was proposed.

Using the AVL model, an addition of a horizontal tail allowed us to estimate the distance of the

horizontal tail in order to stabilize the model. By adding a horizontal tail, the aerodynamic center shifts

further from the nose. There is a consequence of the C.G moving along with the A.C. in which we are

hoping that the A.C moves faster than the C.G.

The table above shows a list of the distance of the horizontal tail with respect to the nose along

with the corresponding A.C and stability derivatives. The highlighted column indicated the location of

the horizontal tail that will separate the AC aft the CG by 1 inches.

With No Tail AC/NP (ft) AC (in) CG (ft) CG (in) CLalpha CYalpha Clalpha Cmalpha Cnalpha AC-CG (ft) AC-CG (in)

0 0.715441 8.585292 0.708333 8.499996 3.343667 -4.9E-05 -0.000032 -0.07202 0.000008 0.007108 0.085296

1.25 0.715441 8.585292 0.708333 8.499996 3.343667 -4.9E-05 -0.000032 -0.07202 0.000008 0.007108 0.085296

Horizontal tail's LE

Distance from the

nose (ft) AC/NP (ft) AC (in) CG (ft) CG (in) CLalpha CYalpha Clalpha Cmalpha Cnalpha AC-CG (ft) AC-CG (in)

1.25 0.765266 9.183192 0.708333 8.499996 3.878691 -3.7E-05 -0.000035 -0.66917 0.000009 0.056933 0.683196

1.5 0.788308 9.459696 0.708333 8.499996 3.832834 -3.8E-05 -0.000033 -0.92888 0.000009 0.079975 0.9597

1.75 0.811366 9.736392 0.708333 8.499996 3.814559 -3.5E-05 -0.000035 -1.19099 0.000009 0.103033 1.236396

2 0.834437 10.013244 0.708333 8.499996 3.806894 -3.3E-05 -0.000034 -1.45474 0.000009 0.126104 1.513248

2.25 0.857456 10.289472 0.708333 8.499996 3.802936 -3.3E-05 -0.000033 -1.7185 0.000009 0.149123 1.789476

Stability

Derivatives

Stability

Derivatives

Table 6-1 Predicted stability derivatives with respect to the horizontal tails distance

47

6.3 THEORY PREDICTIONS OF MODEL STABILITY DERIVATIVES USING

ATHENA VORTEX LATTICE (AVL) The data reduction from the pitch pulse testing provided us the damping and the frequency

response of the system. In order to compare these results to theory we used Athena Vortex Lattice’s

linear theory to provide us with stability derivatives. The derivatives obtained from AVL are then

inputted to the short period approximation equations found on Robert C. Nelson’s book Flight Stability

And Automatic Control. The equation are as follows:

𝜔𝑛𝑆𝑃= √

𝑍𝛼𝑀𝑞

𝑢0− 𝑀𝛼 = 2𝜋𝑓

𝜁𝑆𝑃 = −𝑀𝑞 + 𝑀�̇� +

𝑍𝛼

𝑢0

2𝜔𝑛𝑆𝑃

The longitudinal derivatives on these equations are found on Table 3.5 or Table 4.2 of Nelson’s

book. Some additional information such as estimating the longitudinal stability coefficients are also

found on Table 3.3 of Nelson’s book. For the longitudinal derivatives that contains a coefficient of lift

due to the change of angle of attack, they were set to be 2π for an ideal case. The varying parameter here

is the velocity. Using AVL, coefficient of moment due to change of angle of attack were calculated and

is shown in Table 6.2-1

Table 6-2 Variation of cmα for different velocities using AVL

velocity cmα

70 -1.63701

80 -1.66397

90 -1.67737

100 -1.68445

110 -1.68836

120 -1.69057

With these values damping and frequency were calculated using a Matlab code. These values

were then plotted against the data acquired from wind tunnel testing and Figure 6-1 and 6-2 shows these

comparison.

48

Figure 6-1 Frequency vs. Velocity comparison for theoretical and experimental

Figure 6-2 Damping vs. Velocity comparison for theoretical and experimental

As you observe from the figures, the experimental damping and frequency agrees with the

theoretical values with a few bad data points. These graphs shows how close the model’s theoretical

short period response to a pitch response experiment conducted in real life.

0

0.5

1

1.5

2

2.5

3

70 80 90 100 110 120

Fre

qu

ency

(H

z)

Velocity (ft/s)

Frequency vs. Velocity

Theory

Data

0.15

0.175

0.2

0.225

0.25

0.275

0.3

70 80 90 100 110 120

Dam

pin

g

Velocity (ft/s)

Damping vs. Velocity

Data

Theory

49

7.0 SIMULINK REAL-TIME CONTROL SYSTEM: (SEE

REFERENCE)

7.1 SIMULINK MODEL CONFIGURATION: (SEE REFERENCE)

7.2 THE REAL-TIME WINDOWS TARGET: (SEE REFERENCE)

7.2.1 Advantages of the Real-Time Windows Target (See Reference)

7.2.2 Known Issues with the Real-Time Windows Target (See Reference)

7.3 THE X-56 DART FLIGHT CONTROLS MODEL: (SEE REFERENCE)

7.3.1 The ArduPilot Mega(APM) Interface Subsystem (See Reference)

7.3.2 Pilot Input Subsystem (See Reference)

7.3.3 The System Status Subsystem (See Reference)

7.3.4 The Wind Tunnel Data Recorder Subsystem (See Reference)

50

8.0 WIND TUNNEL DATA ACQUISITION AND ANALYSIS

8.1 WIND TUNNEL TEST DATA ACQUISITION (SEE REFERENCE)

8.2 WIND TEST DATA ANALYSIS METHOD (SEE REFERENCE)

8.2.1 Longitudinal Stability Tests Example Calculations (See Reference)

8.2.2 Gust Response Test Example Calculations The Gust Response Analysis was use in order to find the X-56 Type natural frequency. It is done by

taking peak to peak amplitudes of the gust response data max was plotted vs. the frequency of the dual

gust vanes.

51

9.0 WIND TUNNEL TEST RESULTS

9.1 LONGITUDINAL STABILITY FREQUENCY AND PITCH DAMPING

TEST

9.1.1 Elevator Pulse Excitation Method Results The Elevator Pulse Excitation Method was to simulate a pitch doublet with the X-56 Type model

during trimmed flight within the low speed wind tunnel. The excitation test was use to create a

interference within the models flight to achieve a response. The result of the elevator pulse excitation

test was that the model was unable to reenact a pitch doublet therefore unable to acquire any feedback.

The lag time delay between the control of the joystick to the model and the speed of the servos were

unable to deliver the necessary result aim from this test.

9.1.2 Stick-Hit-Nose Boom Excitation Method Results The pitch damping coefficients found through Equations 6.2-1 and 6.2-2 were averaged at each

velocity and plotted as shown in Figures 9.1.2-1 and 9.1.2-2:

Figure 9-1 Pitch Rate Damping versus Velocity

0.1914

0.1878

0.1982 0.1983

0.2855

0.1936y = 0.001x + 0.1159

0.0000

0.0500

0.1000

0.1500

0.2000

0.2500

0.3000

70 80 90 100 110 120 130

Dam

pin

g C

oef

fici

ent

Velocity (ft/s)

Damping vs Velocity

52

Figure 9-2 Pitch Rate Frequency versus Velocity

The equations shown in the figures show the trend line of the data points. It is found that the pitch

rate damping coefficient trends upwards at higher velocities. This speaks to the model’s pitch becoming

more stable when flying at higher speeds, which matches with previous observations of finding stable

trim of the model being easier as wind speed was increased. Similarly, the frequency increases

according to velocity, proving our theory generated AVL empirical equations on the frequency of the

short period mode response of the aircraft.

9.2 GUST RESPONSE MAGNITUDE VS. GUST VANE DEFLECTION AND

FREQUENCY TEST- GUST RESPONSE MAGNITUDE VS. VANE

FREQUENCY AT VARIOUS TUNNEL VELOCITIES

9.2.1 Test Results Similar to the stick-hit boom excitation testing, the pitch rate damping coefficient for short

period mode was attempted to be determined through testing using a simulated gust. Then the gust vanes

and the model pitch-rate allowed to decay. To accomplish this, wooden gust vanes were installed into

CPP’s subsonic wind tunnel at an angle of ±4.75 degrees. Similar tests were done with the gust vanes

oscillating at different frequencies to replace the manual pitch pulse used in initial testing. Gust vane

pitch response is shown for V=90 ft/s in the figures below:

1.184126984

0.937637509 0.938492063

1.179365079

1.448571429 1.650029198y = 0.014x - 0.1281

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

70 80 90 100 110 120 130

Fre

qu

ency

(cy

/s)

Velocity (ft/s)

Frequency vs Velocity

53

Figure 9-3 Pitch Rate Damping Response with Gust Vane oscillating at 1.25Hz



54




55



9.2.2 Finding Natural Frequency through Gust Response Analysis With the gust vanes continuous oscillations, the maximum pitch rate versus vane pitch frequency

was determined to identify the pitch rate natural frequency versus vane frequency. The maximum

oscillating peak to peak pitch angle due to the gust vane excitation versus gust vane frequency is shown

in Figure 9.2.2-1. The frequency at the maximum amplitude should be the short period natural

frequency. The data recorded was only usable at V = 90 ft/s due to data corruption for other velocities.

56

Figure 9-11 Model Maximum Peak to Peak Oscillating Pitch Amplitude vs. Gust Vane Frequency @ V = 90 ft/s

The trendline set to polynomial of the second order plot shows a natural frequency of 2.6371

cycles/second.

9.3 ERROR AND PROBLEMS This year, the GUst Alleviation and Controls team had encountered errors and problems that

created a bump and challenge within the project. The problems that occurred during the year were:

o Inadequate funding

o Structural Support of the Model (Base)

o No Reliable Power Source

o Broken Servos

o Single Computer Source

o Access to Low Speed Wind Tunnel

o Airpockets

These problems were faced during the year but were overcome as time progress. Within the project

itself, errors that were found were:

o MATLAB/Simulink Failure

o Not Able to reenact pitch doublet

o Right Aileron Servo Failure

y = -1.5921x2 + 8.3972x - 7.5469

0

1

2

3

4

5

6

7

0 1 2 3 4 5

Max

imu

m P

eak

-Pea

k O

scil

lati

ng

Pit

ch

Am

pli

tud

e

Frequency (cycles/sec)

Gust Vane Pitch Amplitude vs. Frequency

57

10.0 CONCLUSION By adding a tail and elevator to the X-56 type model mounted on a sting with a free- free flight

gimbal for pitch and plunge, trim flight was achieved in the Cal Poly Pomona Low Speed Wind Tunnel.

With trimmed flight, an elevator pulse was attempted in order to simulate a small gust disturbance to the

craft, however this gave no discernable results. The vertical excitation with an external rod hitting the

nose boom provided much clearer results and showed stable damping with short period mode, indicating

that the model was indeed capable of stable flight when disturbed or in turbulent flight conditions. The

hit- decay pitch angle and rate data was used to calculate the short period stability frequency and

damping the model. Testing the model in continuous sinusoidal gust field induced by the wind tunnel

oscillating dual gust vanes provided another way of finding the model’s natural frequency and response

magnitude in a gust environment. With all the data receive and testing, Team GUAC was able to

complete and meet various objectives that were set in the beginning of the project term. Future work that

would needed in continuing the research would be seen in Section 11.

58

11.0 RECOMMENDATION Recommendations

Below are the recommendations organized in each area of the X-56 Type Program. These

recommendations can be past down and used by future groups for the benefit of X-56 Type Program.

Structures/Fabrication

Use graphite epoxy for the tail, and elevator

Rebuild the fuselage with high density foam

Purchase a reliable power source of min 5V

Faster servo for the new tail

Replace rigid wings with flexible wings

Mass balance the flexible wings

Controls

Don’t have wires placed within the wings of the model

Color code wires

Re-label the input channels properly

Purchase a new Arduino board that has more input channels

Fix the body flap so the moment arm can be created only with the new elevator

Aerodynamics

Fix the broken servos so all the controls surfaces work on the blended wing model

Funding

Apply early for Cal Poly Grants

Look for additional sources of funding

Testing

Add Gust Alleviation Flight Control Law

Test Gust Alleviation Control Law at the same conditions of base line

59

12.0 CONTRIBUTIONS AND ACKNOWLEDGMENTS The GUst Alleviation and Controls Team (GUAC) for the X-56 Type would like take this

opportunity thank the following for the contribution and part with the project. We would like to thank

Cal Poly Pomona for the ability to use and run the Low Speed Wind Tunnel to conduct the various test

as well as the equipment used to help the project endeavors. We would also like to thank the following

people:

James Ceasari (Jim)

Dr. Ahmadi

Dr. Edberg

Amy Currier

Umbra

AIAA

Ramon (Previous Falcon member).

Brian Kelly

60

13.0 REFERENCES

13.1 PREVIOUS SENIOR PROJECT REPORT (TEAM FALCON BY EVAN

JOHNSON) – LOCATED IN PROFESSOR DOBBS’S LIBRARY

13.2 MATLAT PLOTS

13.2.1 Stick Excitation Test Plots

13.2.1.1 Velocity at 77 fps

64


67


70


73


76


79

13.2.2 Frequency Excitation data


Test Number velocity (fps) vane angle (+/-) Sting angle (deg) VFD Freq. (Hz) VFD RPM Flywheel RPM Flywheel RPS

4.0 0 4.5 0 0 0

4.1 0 4.5 0 16.67 500 50 0.833333333

4.2 0 4.5 0 16.67 500 50 0.833333333

4.3 10 4.5 0 16.67 500 50 0.833333333

4.4 20 4.5 0 16.67 500 50 0.833333333

4.5 30 4.5 0 16.67 500 50 0.833333333

4.6 40 4.5 0 16.67 500 50 0.833333333

4.7 50 4.5 0 16.67 500 50 0.833333333

4.8 60 4.5 0 16.67 500 50 0.833333333

4.9 60 4.5 0 16.67 500 50 0.833333333

80

4.10 70 4.5 0 16.67 500 50 0.833333333

4.11 80 4.5 0 16.67 500 50 0.833333333

4.12 80 4.5

16.67 500 50 0.833333333

4.13 80 4.5

16.67 500 50 0.833333333

4.14 90 4.5

16.67 500 50 0.833333333

4.15 90 4.5

16.67 500 50 0.833333333

4.16 90 4.5

16.67 500 50 0.833333333

4.17 90 4.5

16.67 500 50 0.833333333

4.18 0 4.5

25 750 75 1.25

30.03 900 90 1.5

35 1050 105 1.75

40 1200 120 2

50.03 1500 150 2.5

60.03 1800 180 3

70.03 2100 210 3.5

80 2400 240 4

90.03 2700 270 4.5

100 3000 300 5

81


TEST1 V 100 FPS

AOA 8 deg

Trim Times

Comments

C/S 1.25 1

Not able to get stable flight

T0 0 2 TF

3

4

5

Trim Times

Comments

C/S 1.5 1 2880 T0 2880 2 2912 TF

3 2950

4

5

Trim Times

Comments

C/S 1.75 1 3235 T0 3235 2 3255 TF

3 3265

4

5

Trim Times

Comments

C/S 2 1 3395 T0 3395 2 3400 TF

3 3500

4

82

5

Trim Times

Comments

C/S 2.5 1 3760

may hit bottom stop T0 3760 2 3890

TF

3 3910

4

5

Trim Times

Comments

C/S 3 1 4000

shortest time to achieve trim

T0 4000 2 4015

brian used the force. TF

3 4020

and his ridiculously good loooks

4

5

Trim Times

Comments

C/S 3.5 1 4120 T0 4120 2 4130 TF

3 4170

4

5

Trim Times

Comments

C/S 4 1 4205

getting easier to oscilate in trim with gust vanes

T0 4205 2 4210 TF

3 4215

4

83

5

Trim Times

Comments

C/S 4.5 1 4405

went to 5 on first attempt. Trim times will be later than 5.0

T0 4405 2 4425

ish? TF

3 4445

4 4490

5

Trim Times

Comments

C/S 5 1 4285

did before 4.5 T0 4285 2 4295

TF

3 4325

4

5


TEST

**started from 5.0 and went down to 1.25**

V 110 FPS AOA 8 deg

Trim Times

Comments

C/S 1.25 1

amplitude is larger than gimble range. Says brian. Hes right

T0 0 2

always right TF

3

and cute

4

always cute

5

84

Trim Times

Comments

C/S 1.5 1 5860 T0 5860 2 5870 TF

3 5910

4

5

Trim Times

Comments

C/S 1.75 1 5695 T0 5695 2 5735 TF

3 5785

4

5

Trim Times

Comments

C/S 2 1 5545 T0 5545 2 6505 TF

3 6515

4

5

Trim Times

Comments

C/S 2.5 1 5405 T0 5405 2 5470 TF

3

4

5

Trim Times

Comments

C/S 3 1 5265

difficult to attain stable trim for more than 1 second.

85

T0 5265 2 TF

3

4

5

Trim Times

Comments

C/S 3.5 1 4920 T0 4920 2 4925 TF

3 4945

4 4960

5

Trim Times

Comments

C/S 4 1 4840 T0 4840 2 4845 TF

3 4855

4

5

Trim Times

Comments

C/S 4.5 1 4700 T0 4700 2 4710 TF

3 4740

4

5

Trim Times

Comments

C/S 5 1 4620 T0 4620 2 4650 TF

3 4630

4

86

13.3 EXCEL DATA RESULTS

13.3.1 Stick Excitation Test


Hit 1 0.19576058

77

2 0.395500034

77 3 0.174720297

77

4 0.206965088

77 5 0.188089312

77

6 0.095047332

77

AVERAGE 0.209347107

CorAVG 0.191383819


Hit 1 0.39667778

80

2 0.196524524

80 3 0.174592973

80

4 0.109880573

80

5 Omitted Lost Trim 80

6 0.192354664

80 AVERAGE 0.214006103

CorAVG 0.187824054


Hit 1 0.231888752

90

2 0.220620582

90 3 0.112013249

90

4 Omitted Lost Trim 90

5 0.176280015

90 6 0.163865929

90

AVERAGE 0.180933706

87

CorAVG 0.19816382


Hit 1 0.271327347

100

2 0.252005937

100 3 0.186688748

100

4 0.19318227

100 5 0.161160295

100

6 0.107938882

100 AVERAGE 0.195383913

CorAVG 0.198259313


Hit 1 0.381239751

110

2 0.594857515

110 3 0.257067064

110

4 0.256143965

110 5 0.247485937

110

6 Omitted

110 AVERAGE 0.347358846

CorAVG 0.285484179


Hit 1 0.193118608

120

2 0.220620582

120 3 0.354660875

120

4 0.123886208

120 5 0.059078315

120

6 0.236790724

120 AVERAGE 0.198025885

CorAVG 0.193604031

88

13.4 BOOKS

Nelson, Robert. Flight Stability and Automatic Control. 2th

ed. McGraw-Hill Book Co 1998

13.5 OTHER DOCUMENTS

Previous Years Final Report (Located in Professor Dobb’s Library)

89

13.6 POSTER COPY

Gust Alleviation & Con trol Systems Of An X-56 Type Aircraft

Wi nd T unnel Model

Objective: - Achieve A Successful Trim Flight With A Horizontal Tail.

- Obtain And Calculate Stability Derivative For Short Period Mode

- Measure Gust Response To Locate Aircrafts Natural Frequency

!

! "#$"%&'( %) %$*"'!"#$%!&'%( !)*!

+, - , .!/0- !1!0230!455'%( 6 78#9: ; : %$<=5#!

+*, - ./'! "#$"%&'( %) %$*"!&> '?( 4!@$A$!

+0B3.!C- - !1!- 3, 2!5> '?( 4%$A$6 ?; $'D<7: ; !

0*%1'2#) ."#3'4/5.*&5'6) $7) **"'E =: *?=!F$?#': !+0B3.!3- C!1!2- 3G!

??9$?#': 6 78#9: ; : %$<=5#!

2#) ."#3'4/5.*&5'6) $7) **"!&: %?!)'%!HI : : !+/BC.!J /3!1!02/0!

K#8L%<*I : : 6 ?; $'D<7: ; !

0*%1'8*"#1/) %&795'6) $7) **"!M'DD!N : ?$( %!

+, - , .!, , 3!1!0J , / !> : ?$%#O6 I $( : : <7: ; !

8*"#1/) %&795'6) $7) **"!H=?'%$D5!E #'%4: !+, - , .!23, !1!2- , - !

*=?'%$D5?#'%4: 6 ?; $'D<7: ; !

0*%1'4."- 9.- "%3'6) $7) **"!"$%%=*!PD$*Q!

+GGB.!0B, !1!3, C3!477D$*Q6 78#9: ; : %$<=5#!

2: 7*;'<%="79%>#) '6) $7) **"!)$8: %!R: %?!

+G0G.!GG- !1!/G- 3!K$8: %Q: %?6 78#9: ; : %$<=5#!

2: 7*;'<7) %) 97%3'? @9*"!S*I $!N 'DD'$; 8!

+G0G.!/B- !1!- 0B, !$I *$K#%=> 'DD'$; 86 ; 8%<7: ; !

8*"#1/) %&795'4- , , #". !P*I 84$D!@#%=T!

+/BC.!GB, !1!BC3C!7*I 84$D<%#%=T3J 6 I $( : : <7: ; !

8*"#1/) %&795'4- , , #". !S%$( '!U=*%$%5=T!+303.!C, - !1!0, 2, !

$%$( '( 6 78#9: ; : %$<=5#!

! "#$"%&'%) 1'2#) ."#3'4/5.*&5'81A75#"!!VA$%!): ( %8: %!+/BC.!J 2B!1!CBCG!

=*K: ( %8: %00/6 ?; $'D<7: ; !

<%9- 3./'81A75#"!W4=A=%!&: OO8!

Method:- Add a horizontal tail with elevator control to X-56 type model

- Attach plane to free ight gimbal mount in order to test for trim ight

- Fix gimbal mount to make it more compatible with the new horizontal

tail

- Test at varying speeds and angle of attack to achieve trim ight

- Cover base of the plane with tape to prevent “parachuting e ect” while trying to

achieve trim ight

- Perform an elevator pulse-doublet in order to test for short period stability

- Test for short period stability by exciting the craft externally should elevator pulse

fail to produce results

- Install gust vanes in the Cal Poly Low Speed Wind Tunnel

- Vary gust vane frequency from .5 Hz to 3 Hz

Recommendation:These recommendations can be past

down and used by future groups for the

bene t of X-56 program.

Structures/Fabrication

• Better materials for the tail, and

elevator

• Rebuild the fuselage with high

density foam

• Purchase a reliable power source of

min 5V

• Faster servo for the new tail

Controls

• Don’t have wires placed within the

wings of the model

• Color code wires

• Re-label the input channels properly

• Purchase a new Arduino board that

has more input channels

• Fix the body fla

p

so the mo me n t arm

can be created only with the new

elevator

Aerodynamics

• Fix the malifunctioning servos to all

the controls surfaces work on the

blended wing model

Conclusion: By adding a tail and elevator to the x-56 type model, trim flig h t

was achieved. With trimmed flig h t, an elevator pulse was attempted

in order to simulate a small gust disturbance to the craft, however

this gave no discernable results. The excitation with an external rod

on the other had provided much clearer results, indicating that the

model was indeed capable of stable ight when disturbed or in tur-

bulent ight conditions. The data we were able to obtain from the

experiment was used to calculate the short period stability of the

model. Adding the gust vanes provided another way of nding the

model’s natural frequency.

Approach:- Approximate X-56 type model CG and AC locations. Design tail that

will move AC further back while not a ecting CG location.

- Create Horizontal Tail to add to Existing X-56 Model

- Add functioning elevator to control pitch

- Lock wingtip rudders

- Use Cal Poly Low Speed Wind Tunnel to achieve stable ight

- Attach model to gimbal mount, then attach model/gimbal mount apparatus to crescent

sting in order to maintain free ight

- Use Team Falcon’s Simulink model in order to control model with joystick

- Use the on board gyros and accelerometer to measure pitch, pitch rate, and time

- Simulate turbulent wind conditions by uctuating pitch. This can be accomplished by

performing an elevator pulse-doublet as well as by tapping the nose boom with a rod.

- Vary gust vanes in tunnel in order to nd the optimal speed needed to

fin

d

the mo del ’s natural frequency

Data:

- AVL Model

Gust Vane ResponseStick Test Reponse

Team GUAC-

Figure 13. 6-1 Team GUAC Poster Board

90

APPENDIX A-1: CODE LISTING

A-1.1 MATLAB CODE FOR SHORT PERIOD APPROXIMATION clc; clear; %Aircraft geometry and mass data cbar = 0.33; %Mean aerodynaic chord (ft) b = 2.775; %Wing span (ft) S= 1.11; %Wing planform area (ft^2) St = (8*2+1.5*7.25+1.5*0.375)/144 %Horizontal tail planform area (ft^2) lt = 0.875; %Distance of the 1/4 chord of the tail

to the C.G. (ft) AR = b^2/S; %Aspect Ratio of the wing W = 4.27; %Weight of the aircraft (lbs) gtos = 0.000068521765562; Iyy = (190*gtos)*(17.375/12)^2+(1626*gtos)*(1.375/12)^2+(122*gtos)*(9.75/12)^2;

%Moment of inertia about y-axis (slug ft^2) g = 32.2; %Acceleration due to gravity (ft/s^2) m = W/g; %Mass of the Aircraft (slugs) %Flight Condition Data V = 120; %Trim Speed (ft/sec) u0 = V; rho = 0.002378 %Flight Density (slugs/ft^3) Q = 0.5*rho*u0^2; %Flight Dynamic Pressure at Trim

(lbs/ft^2)

Cmalpha = -1.690573;

CLalphatail = 2*pi(); CLalphawing = CLalphatail;

dedalpha = (2*CLalphawing)/(pi()*AR) VH = (lt*St)/(S*cbar)

Czalphadot = -2*CLalphatail*VH*dedalpha; Cmalphadot = -2*CLalphatail*VH*(lt/cbar)*dedalpha; Cmq = -2*CLalphatail*VH*(lt/cbar) Zwdot = -Czalphadot*(cbar/(2*u0))*((Q*S)/(u0*m)) Zalpha = u0*Zwdot

Mq = Cmq*(cbar/(2*u0))*((Q*S*cbar)/Iyy)

Mw = Cmalpha*((Q*S*cbar)/(u0*Iyy)) Malpha = u0*Mw

Mwdot = Cmalphadot*(cbar/(2*u0))*((Q*S*cbar)/(u0*Iyy)) Malphadot = u0*Mwdot

wn = sqrt((Zalpha*Mq/u0)-Malpha)

91

zeta = -(Mq+Malphadot+(Zalpha/u0))/(2*wn)

f=wn/(2*pi()) g=zeta*(2*pi())

APPENDIX A-2: AVL INPUT FILES

A-2.1 GEOMETRY FILE (.AVL) GUst Alleviation and Controls

#UNITS ARE IN FEET

#Mach

0.0

#IYsym IZsym Zsym

0.0 0.0 0.0

#Sref Cref Bref

1.11 0.33 2.775

#Xref Yref Zref

0.677083 0.0 0.0

#1.0 Fuselage====================================================

SURFACE

Fuselage

#Nchordwise Cspace Nspanwise Sspace

10.0 1.0 80.0 0.0

YDUPLICATE

0.0

ANGLE

0.0

#Center Line--------------------------------------------------

SECTION

#Xle Yle Zle Chord Ainc Nspanwise Sspace

0.0 0.0 0.0 1.2330 0.0 0.0 0.0

AFILE

Fuselage.dat

#CONTROL

#Cname Cgain Xhinge HingeVec SgnDup

#ELEVATOR 1.0 0.8737 0.0 1.0 0.0 1.0

#1.2--------------------------------------------------

SECTION


0.2160 0.1249 0.0 1.0172 0.0 0.0 0.0

AFILE

Fuselage.dat

#CONTROL


#ELEVATOR 1.0 0.8737 0.0 1.0 0.0 1.0

#1.3--------------------------------------------------

SECTION


0.3030 0.1690 0.0 0.9301 0.0 0.0 0.0

AFILE

92

Fuselage.dat

#CONTROL


#ELEVATOR 1.0 0.8737 0.0 1.0 0.0 1.0

#1.4--------------------------------------------------

SECTION


0.4379 0.2346 0.0 0.5525 0.0 0.0 0.0

AFILE

Fuselage.dat

#CONTROL


#ELEVATOR 1.0 0.8737 0.0 1.0 0.0 1.0

#1.5--------------------------------------------------

SECTION


0.5273 0.3125 0.0 0.4252 0.0 0.0 0.0

AFILE

Fuselage.dat

#CONTROL


#ELEVATOR 1.0 0.8737 0.0 1.0 0.0 1.0

#2.0 Wing====================================================

#2.1 (ROOT)--------------------------------------------------

SECTION


0.6521 0.5412 0.0 0.2450 0.0 0.0 0.0

AFILE

Wing.dat

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A-2.2 RUN CASE FILE Run case 1: Trim

alpha -> alpha = 0.00000

beta -> beta = 0.00000

pb/2V -> pb/2V = 0.00000

qc/2V -> qc/2V = 0.00000

rb/2V -> rb/2V = 0.00000

alpha = 5.00000 deg

beta = 0.00000 deg

pb/2V = 0.00000

qc/2V = 0.00000

rb/2V = 0.00000

CL = 0.463351

CDo = 0.00000

bank = 0.00000 deg

elevation = 0.00000 deg

heading = 0.00000 deg

Mach = 0.00000

velocity = 73.3300 Lunit/Tunit

density = 0.237700E-02 Munit/Lunit^3

grav.acc. = 32.2000 Lunit/Tunit^2

turn_rad. = 0.00000 Lunit

load_fac. = 1.00000

X_cg = 0.643900 Lunit

Y_cg = 0.00000 Lunit

Z_cg = 0.00000 Lunit

mass = 3.75000 Munit

Ixx = 1.11267 Munit-Lunit^2

Iyy = 1.46404 Munit-Lunit^2

Izz = 2.54078 Munit-Lunit^2

Ixy = 0.00000 Munit-Lunit^2

Iyz = 0.00000 Munit-Lunit^2

Izx = 0.00000 Munit-Lunit^2

visc CL_a = 0.00000

visc CL_u = 0.00000

visc CM_a = 0.00000

visc CM_u = 0.00000

96

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102

APPENDIX A-3: PROPOSAL

Senior Project Proposal

TEAM:

GUst Alleviation Control (2014-15)

TEAM MEMBERS:

Tuan Dinh Jr [email protected] (909) 720-2532 Team Lead

Dwight Nava [email protected] (213) 400-0395 Deputy

Reginald Guinto [email protected] (909) 539-5090

George Paguio [email protected] (213) 304-5036

Tanner Clark [email protected] (661) 219-3943

Jason Kong [email protected] (626) 660-7603

Dong Jin Ryoo [email protected] (714) 873-2572

Arya WIlliams [email protected] (626) 710-0219

Bill Wogahn [email protected] (909) 993-2897

CONSULTANT AND SUPPORT TEAM:

Evan Robert Johnson [email protected] (714) 851-4146

Crystal Nunez [email protected] (714) 619-1434

Anahi Hernandez [email protected] (323) 490-2959

Faculty Advisor: Professor Steven K. Dobbs

Submission Date: October 2014

Approved: ______________________________________________, Date _________________

AEROSPACE ENGINEERING DEPARTMENT

CALIFORNIA POLYTECHNIC UNIVERSITY, POMONA

103

1.0 NEED AND BENEFITS OF THE PROJECT:

1.1 Next generation aircrafts are implementing higher AR wings (Blended Wing Bodies). Further

research on aero-elastic phenomena, specifically the analysis of flutter and gust response

challenges, will contribute to creating successful next generation aircrafts.

1.2 Of the four major aerospace engineering disciplines, control systems offers the least

undergraduate courses despite having become one of the most important disciplines in industry.

This project allows for a greater understanding of advanced control topics not covered in any

undergraduate course.

1.3 If the department were to include practices and test procedures from this project in the

aerospace curriculum in the form of a lab, undergraduate students would be exposed to data

acquisition and controls hardware.

1.4 Students will be provided with hands on experience with testing a blended wing-body’s

Stability Augmentation System (SAS).

2.0 PRIMARY PROJECT OBJECTIVES:

2.1 Modifying the blended wing body model fabricated by the 2013-2014 FALCON Club senior

project team. The modified X-56 will have a relocated C.G by the use of a nose boom weight in

front of the model. The goal of this is to move the AC aft of the C.G. for static and dynamic

stability.

2.2 The FALCON (Flutter ALleviation and CONtrol) model will be tested in the subsonic wind

tunnel with two degrees of freedom for longitudinal stability testing in order to develop an

optimized longitudinal Stability Augmentation control system.

2.3 Demonstrate a longitudinal gust alleviation system capable of reacting to vertical gusts in the

subsonic wind tunnel. This test will utilize the gust generation system designed and installed in

the wind tunnel by the 2012-2013 Flutter Club team.

2.4 Expand stability augmentation system to include lateral-directional motion. Demonstrate

controllability and augmented static and dynamic stability of a blended wing-body aircraft in

five degree of freedom motion.

104

2.5 Super impose the stability augmentation system with gust alleviation system to reduce flutter

phenomena for a rigid wing structure.

2.6 Perform preliminary vibration and flutter analysis in NASTRAN for a composite-skinned

flexible wing optimized for span wise torsional bending. Determine optimal material, wing

structure, and mass distribution to obtain desired structural dynamic modes.

3.0 SECONDARY PROJECT OBJECTIVES:

3.1 Explore and test alternative control methods for a blended wing-body aircraft. Utilize

differential rudder and aileron actuation in order to aid in yawing motion through differential

drag effects.

3.2 Demonstrate the feasibility of a remotely processed control system running on a portable

computer running a real-time Simulink control system.

3.3 Provide the Aerospace Engineering department with a fully functional flutter research model

capable of modeling advanced control concepts aimed at actively suppressing aero-elastic flutter

during simulated flight in the low-speed wind tunnel.

4.0 PRODUCTS OF PROJECT:

4.1 New updates to the FALCON platform including a variable C.G. through the use of a nose

boom weight, as well as a horizontal tail with elevator and vertical tails with rudder for an

increase in control surface sensitivity.

4.2 Valuable research for real time stabilization of an unstable blended wing body aircraft. Included

in this research is the capability of stabilizing an aircraft that uses flexible, lightweight

composite wings.

5.0 METHODS TO MEET PROJECT

105

OBJECTIVES:

5.1 Laboratory Use Required:

Aerospace Engineering Department Low Speed Wind Tunnel

Structures Laboratory

UAV Laboratory

Thermal Fluids Laboratory

5.2 Experiments and Tests:

5.2.1 Stable static margin

Determine location of the C.G. with respect to the nose boom weight.

Hand calculation of expected aerodynamic center.

Graphical representation between C.G. and nose boom weight location with respect to

the calculated A.C.

5.2.2 Flight Testing:

Implement a tail (if necessary) to create a bigger moment arm for more effective

aerodynamic control of trim and stability.

Develop a longitudinal Stability Augmentation System.

Introduce a gust alleviation system using the SAS.

5.3 Equipment Needed for Project:

ArduPilot Mega 2.5 integrated flight management system

High speed ADC (Analog to Digital) equipment.

Wire resistance strain gauges.

Accelerometer/Gyroscope packages. (Inertial Measurement Units)

Low-speed wind tunnel 5-DOF gimbal mount.

Vibrational analysis equipment:

OROS Dynamic Analysis Hammer

5.4 Computer Software Required:

Solid Works 3D modeling and analysis software.

Athena Vortex Lattice

NASTRAN structural analysis software.

MATLAB control simulation software:

SIMULINK Real-Time Libraries

SIMULINK Code Generation Packages

106

5.5 Test Articles:

Existing FALCON X-56 model.

Carbon Fiber/Composite Resin horizontal tail for X-56 FALCON.

Fully customizable x86 integrated flight management and control system for FALCON

system.

5.6 Cost Breakdown:

107

5.7 Sources of Funding

Program Who is eligible

to apply

Purpose Funds for

Student

Funds For

Faculty

Deadlines

Engineering

Scholars

Program (ESP)

Engineering

majors, Pell

Grant eligible

Conduct UR and

prepare for

graduate school

and career

development

$5,000 per

Academic Year

N/A October 14th

Spring 2015

Recruitment for

continuing

students for

2015-16

Academic Year

President’s

Special Travel

Fund

ALL students,

All colleges all

majors

Travels to

present at

conferences

Up to $1,000

(registration,

Transportation,

Hotel)

Conference Fee

Registration

October 20th

Kellog

Undergraduates

Scholar Program

Low income, FG,

and military

veterans. GPA of

2.7 or higher

Faculty mentored

research projects

Up to $2000 per

Academic Year

See Kellog

Future Mini

Grant

November 6th

Kellog Future

Mini-Grants

Faculty and their

student research

treams

Minor

equipment,

software,

research travel

At discretion of

faculty mentor

Up to 2000 per

academic year

November 6th

https://www.csupomona.edu/~our-cpp/student-resources/engineering-scholars-program.shtml



108

6.0 TEAM ORGANIZATION CHART

Program Manager

Tuan Dinh Jr

Lead Control Systems Engineer

George Paguio

Control Systems Engineer

Dong Ryoo

Lead Aerodynamics Engineer

Bill Wogahn

Aerodynamics Engineer

Reginald Guinto

Aerodynamics Support

Crystal Nunez

Aerodynamics Support

Anahi Hernandez

Lead Structural Engineer

Tanner Clark

Structural Engineer

Jason Kong

Chief Fabrication Engineer

Jason Kong

Fabrication Engineer

Tanner Clark

Chief Financial Officer

Arya Williams

Deputy Program Manager

Dwight Nava

Program and Control Systems Advisor

Evan Johnson

Faculty Advisor

Prof. Dobbs

109

7.0 KEY RISKS TO MEETING OBJECTIVES –

MITIGATION PLAN 7.2 Horizontal tail structural failure

Do structural test for the composite rods

Lay up the plywood with a composite skin

Trade studies on materials

7.3 Inadequate funding

Come up with additional funding options.

If necessary fund project costs out of pocket

7.4 Inadequate training in autonomous stability control systems

Contact Evan Johnson for a briefing about autonomous stability control systems

Do research on autonomous stability control systems

8.0 SCHEDULE OF MAJOR PROJECT MILESTONES

Name Start Finish % Complete

Assemble Team Sep 30 2014 Oct 10 2014 100

Proposal Due Sep 30 2014 Oct 30 2014 100

Static Margin Test Oct 16 2014 Oct 27 2014 50

Perform structural modifications Oct 17 2014 Nov 11 2014

Wind tunnel test #1 w/ Tail (open loop) Nov 10 2014 Nov 24 2014

Fall Quarter Interim Report Oct 16 2014 Dec 1 2014

Develop optimized longitudinal Stability

Augmentation System Oct 13 2014 Jan 31 2015

Wind tunnel test #2 w/ Tail (closed loop “S.A.S”) Feb 1 2015 Feb 28 2015

Winter Quarter Interim Report Jan 5 2015 Mar 13 2015

Develop longitudinal Gust Alleviation System Mar 23 2015 Apr 1 2015

Wind tunnel test #3 w/ Tail (closed loop “G.A.S”)

in Gust flow Apr 1 2015 May 1 2015

Final Report Draft May 1 2015 May 27 2015

Final Report Approved June 1 2015

110

APPENDIX A-4: STICK EXCITATION TEST PLAN 1. Objective:

1.1. Achieve a successful trim flight with a horizontal tail.

1.2. Develop an optimized longitudinal Stability Augmentation control system.

1.3. Achieve data to create a graph of pitch damping vs. speed.

1.4. Achieve data to create a graph of frequency vs. speed.

2. Approach:

2.1. Use the on board gyros and accelerometer to measure angular rotations and

accelerations. Also record slo-mo movies of the side view of the flying model with a

timer (accurate to at least 0.1 seconds) visible in the camera field of vision for frame by

frame measurement of the amplitude vs. time for manual plotting in Excel.

2.2. To simulate a elevator pulse by hitting the nose with a rod

2.3. the control stick, perform an elevator pulse-doublet to excite the model in pitch to

measure the amplitude vs time.

3. Requirements:

3.1. Record digital time

3.2. Measure tunnel air velocity

3.3. Record at every degree until attaining free flight in the middle of the gimbal travel or

reaching 15 degrees angle of attack

3.4. Record decay

3.5. Rod to create elevator pulse for model

4. Method

4.1. Low Speed Wind Tunnel with crescent sting installed to acquire wind tunnel data.

111

4.2. Slowly increase the wind tunnel speed to first test velocity, then gradually increase and

repeat tests up to the maximum velocity.

5. Procedure

5.1. Bench Test

5.1.1. Make sure all the wiring harnesses are connected properly.

5.1.2. Check to see if all the control surface function properly.

6. Flight Test

6.1. Zero out all instruments before initial flight

6.2. Turn on instruments used as well as program to conduct initial testing in Wind Tunnel

6.3. Set Speed and AoA to test matrix requirement

6.4. Change sting AoA to next test matrix value , try to re-trim elevator and repeat increasing

the sting AoA until the sting and elevator trim create free flight in the middle of the

gimbal travel.

6.5. Record Speed and AoA

6.6. Using the control stick, perform an elevator pulse-doublet to excite the model in pitch to

measure the amplitude vs time using a slo-mo movie with a time display and the

Arduino

6.7. Using tunnel-ceiling slot, lightly tap the model to create a pulse sensation.

6.8. Record the decay using a slo-mo movie with a time display (0.1 second accuracy) and

with Pitch rate gyro on Arduino

6.9. Repeat 5.2.5 (2 times)

6.10. Repeat 5.2.3-5.2.7 at a higher speed (5 times)

112

Test Matrix

Test 1

Initial V= 70 ft/sec

Initial Sting AoA= 0

Run #

Sting AoA

(degrees)

V

(ft./s)

Pitch

(degrees) Notes

1.1 0 70

1.2 1 70

1.3 2 70

1.4 3 70

1.5 4 70

1.6 5 70

1.7 6 70

1.8 7 70

1.9 8 70

2 9 70

2.1 10 70

2.2 11 70

2.3 12 70

2.4 13 70

2.5 14 70

2.6 15 70

113

Test 2

Initial V=85 ft/sec

Initial Sting AoA=0

Run #

Sting AoA

(degrees)

V

(Ft/s)

Pitch

(degrees) Notes

1.1 0 85

1.2 1 85

1.3 2 85

1.4 3 85

1.5 4 85

1.6 5 85

1.7 6 85

1.8 7 85

1.9 8 85

2 9 85

2.1 10 85

2.2 11 85

2.3 12 85

2.4 13 85

2.5 14 85

2.6 15 85

114

Test 3


Initial AoA=

Run #

Sting AoA

(degrees)

V

(Ft/s)

Pitch

(degrees) Notes

1.1 0 100

1.2 1 100

1.3 2 100

1.4 3 100

1.5 4 100

1.6 5 100

1.7 6 100

1.8 7 100

1.9 8 100

2 9 100

2.1 10 100

2.2 11 100

2.3 12 100

2.4 13 100

2.5 14 100

2.6 15 100

115

Test 4


Initial AoA=

Run #

Sting AoA

(degrees)

V

(Ft./s)

Pitch

(degrees) Notes

1.1 0 110

1.2 1 110

1.3 2 110

1.4 3 110

1.5 4 110

1.6 5 110

1.7 6 110

1.8 7 110

1.9 8 110

2 9 110

2.1 10 110

2.2 11 110

2.3 12 110

2.4 13 110

2.5 14 110