2007-2008 RIT MAVMid-Term Review

Michael Reeder – Team LeaderKevin Hand – Lead EngineerTodd Fernandez - MESusan Bieck – MEJeremy Teets – MECody Rorick – MEAdam Bosen – CE

…where the sky is only the beginning……and the ground is likely the end…

Put a face to the name…

Mike ReederKevin Hand

Sue Bieck

Jeremy Teets

Adam Bosen

Todd Fernandez

Cody Rorick

Presentation Overview

Organizational Structure Overview of past MAV Projects Introduction/Objective of 2007-2008 RIT MAV Senior Design I deliverables Concept generation Preliminary analysis Platform design and structures analysis Propulsion system selection Airfoil selection MAV control system Design of experiments Indication of progress on deliverable completion

Organizational Structure

Mike Reeder-Team Leader

Kevin Hand-Lead Engineer:

Design of Experiments,

Systems Integration

Todd Fernandez-Propulsion and

Composites

Sue Bieck - Airfoil Analysis

and Aero Structures

Jeremy Teets -CAD Generation

and Aero Structures

Cody Rorick - Flight Dynamics and Component

Integration

Adam Bosen -Component

Integration and Operating Software

Overview of Past MAV Projects

Objectives: Fly 600 m Capture an image Obtain a reliability of 80%

Project went through 5 phases Phase 1: Previous year’s MAV

Firefly motor 300 mA-h battery 6” prop

Phase 2: New Propulsion System 3 x 2” prop Feigao Motor produced more thrust

Phase 3: Angle of Attack (α) Phase 4: Limiting Control Surfaces

Control surfaces determined to be a source of failure

Enabled turning but with limited success

Phase 5: Rudder Design (stiff wing) Throttle use mimics elevator Controlled turning in yaw reduced pilot error 6.5” platform (longest linear dimension)

Phase 5 ResultsBattery, α Flight Time (seconds) Repeatability

No α, 730 mA-h 50 20/100No α, 480 mA-h 117 50/1005° α, 480 mA-h TBD 85/100

Introduction/Objectives of 2007-2008 RIT MAV

Primary Objective: Create a Micro Aerial Vehicle, expandable in nature for future

RIT research, that is simple, robust and stable in design and is capable of reading back information regarding the vehicle’s speed, angle of attack, pitch, yaw and roll rates. Flight Dynamics competition (held internationally) establishes target specifications (engineering metrics) Max linear dimension is 80 cm Max weight is 1 kg Required flight time is 4 minutes

Secondary Objective: Compete in international Flight Dynamics competition

Introduction/Objectives of 2007-2008 RIT MAV

Previous MAVs have been designed around control by an operator with a RC controller on the ground

Control systems are essential to achieving fully autonomous flight

The 2007-2008 RIT MAV will be at a level of development between being fully remote controlled and semi-autonomous with the introduction of control systems being the next step in the development process

Fully Remote Controlled

Introduction of ControlSystems

Fully Autonomous Flight Achieved

2007-2008 RIT MAV autonomous

autonomous

Flight Information

Microcontroller Tri-axial accelerometers

Differential pressure sensors (AOA, velocity)

How do we achieve… …simplicity?

The best airfoil for the application is chosen Fuselage/pod designed around component sizes

…robustness? Foam wrapped in carbon fiber and fiberglass achieves robustness Components used to be placed inside of fuselage/pod for protection

…stability? Minimizing size of platform no longer a concern; therefore, design plane having

a large wingspan with the use of winglets to improve lift Use of recommended aspect ratio in conjunction with rear rudder, ailerons and

elevators supported by “basic” flight dynamic calculations …expandability?

Use of large platform allows for future optimization Microcontroller increases capabilities for future research (see following slides)

Break down of MAV into subfunctions

MAV platform consists of Airfoil and aero structures design

XFOIL analysis Airfoil creation and flight testing “Basic” flight dynamics calculations

Propulsion system Motor selection Prop selection Battery selection

Components/electronics Component selection Design of experiments Preliminary component testing Systems integration

Breakdown of MAV into subfunctions

Semi-autonomous

flight

Flight Information

Microcontroller Tri-axial accelerometers

Differential pressure sensors (AOA, velocity)

Controls

Base station (laptop

configuration)RC controller GPS (future)

Senior Design I Deliverables

Platform design decided upon Engineering metrics/product specifications completed List of components and materials compiled 3-D CAD model of plane created (XFOIL, Pro-E, etc.) Foam model built based on concept generations Experiments designed to test components’ proper functionality Components ordered/in team’s possession Components in possession are in test process Foam plane is built and glide tested

Concept Generation

Started with brainstorming sessions to create various platform and propulsion ideas

Use of Pugh matrices and +, 0, - technique to narrow down spectrum of choices

Team members generated concept drawings/models on their own to emulate narrowed down choices for purposes of visualizing different platforms

Concept Generation: Selection Criteria

A BDucted Fans Propellers

Durability +Thrust +Weight -Cost -Lift Effect -Drag -Size 0Motor Cooling +Sum +'s 3Sum 0's 1Sum -'s 4Rank -1Proceed? No Yes

Selection Criteria

Concepts: Propulsion System

Refe

renc

e

A B D E FFixed Wing Flying Wing Bi-Plane Helicopter Osprey

Expandability - - + +Stable in pitch - 0 0 0Stable in yaw - 0 0 0Stable in roll - 0 0 0High speed + 0 - -Low speed 0 0 + +Weight 0 - - -Visual impact 0 - + +Hand-launchable 0 0 + +Materials needed 0 - - -Advertising space - + - -Simplicity - - - -Sum +'s 1 1 4 4Sum 0's 5 6 3 3Sum -'s 6 5 5 5Rank -5 -4 -1 -1Proceed? Yes No No No No

Selection Criteria

Refe

renc

e

Concepts: Platform Design

**Results: Fixed wing design with propellers for propulsion

Concept Generation: Preliminary Sketches/First Draft Models

Concept Generation: Bill of Materials

After plane concepts were generated, a bill of materials was created reflecting the needs of the MAV

Material/Component Quantity Needed Per PlanePressure Transducer 1Differential Pressure Sensor 2Microcontroller 1Camera 1Accelerometers (tri-axial configuration) 2Batteries TBDMotors 1Speed Controllers 2Propellers 1Servos 4Transmitter/Receiver 1Transceiver (Equivalent) 1Radio Crystals TBDFlexible and Rigid Resin TBDStructural Foam (thicker needed) TBDLEDs 1Watch Batteries 1Wires N/AAntenna N/ABalsa Wood TBDFiber Glass TBDPeel Ply ~ 10 square yardsVacuum Bagging Material ~ 10 square yardsCarbon Fiber ~ 10 square yardsMagnets (access doors) TBD

Preliminary Analysis

Preliminary Analysis

Creation of a stable platform required use of “basic” flight dynamic calculations

Calculations performed produced spectrum of empirically derived formulae to determine stability of platform given certain aspects of the plane design

Preliminary Analysis: Stability

Longitudinal Static Stability Criteria

Directional Static Stability Criteria

Lateral Static Stability Criteria

0m

C & 00mC

0n

C

0l

C

d

dCC mm

L

mm dC

dCC

0

d

dCC nn

d

dCC ll

Where

= Coefficient of Pitching Moment

= Coefficient of Yawing Moment

= Coefficient of Rolling Moment

= Coefficient of Lift

= Angle of Attack

= Sideslip Angle

mC

nC

lC

LC

Preliminary Analysis: Longitudinal Static Stability

Setting to zero and solving for lHSH you arrive at:

lHSH = f( )

Where

= Downwash at 0 angle of attack.

it = Tail incidence angle. iw = Wing incidence angle.

= Coefficient of lift with respect to of the horizontal tail.

= Coefficient of lift with respect to .

Xcg = Center of gravity position along the x axis.

Xac = Center of pressure position along the x axis.

= Coefficient of pitching moment of the fuselage with respect to .

= Coefficient of pitching moment of the wing at 0 angle of attack.

= Coefficient of pitching moment of the fuselage at 0 angle of attack.

fwfwHmmtwmaccgwLLww CCiiCXXARCCSbc00

,,,,,,,,,,,,, 0

0

HLC

wLC

fmC

fmC 0

wmC 0

mC &

0mC

Preliminary Analysis: Directional Static Stability

Setting and solving for lH you arrive at:

lH = f( )

Where Kn = W-B interference factor.

KRl = Correction factor. Sfs = Projected side area of the fuselage.

lf = Length of Fuselage. d = Maximum fuselage depth

Zw = Distance parallel to the z-axis, from wing root quarter chord to fuselage centerline.

= Coefficient of lift with respect to alpha of the horizontal tail.

0n

C

wwLvffsRlnW ARdzCSlSkkSbV

,,,,,,,,,,

VLC

Preliminary Analysis:Lateral Stability

Setting and solving for lH you arrive at:

lH = f( )

Where bH = Wing Span.

= Coefficient of lift with respect to beta of the horizontal tail.

= Coefficient of lift with respect to beta of the vertical tail.

= Coefficient of roll with respect to beta of the wing & body.

wbVHlwwLLVwHH CdzARCCSSSb

,,,,,,,,,

HLC

wblC

lC

0l

C

Platform Design

Platform Design

Initial structural choices were between a Flying wing and a conventional airframe

Due to requirements of the project, a conventional airframe was chosen (see Pugh matrices)

Wing position was chosen to be a top wing configuration for flight stability

With the deciding factor of propulsion being made, additional concept drawings were created

Two designs were generated: a simple design with a tube as a fuselage and another with a tube and blended wing portions

The blending of the wing was done to reduce turbulence effects at the wing-fuselage intersection

Platform Design: Additional Sketches

Platform Design:Prototype 1 – Pro-E Generated

20 inch fuselage 3 inch diameter on the fuselage at largest point 29 ½ inch wingspan Span on horizontal tail is 8 inches Symmetrical airfoil on horizontal tail ¼ inch thick Vertical tail is 4 inches tall

Platform Design: Prototype 1 – Pro-E Generated

Structures Analysis

Structures Analysis:Design Constraints

Minimum weight Maximum durability (crash worthiness)

Withstand estimated (and repeated) 5g crash Withstand belly landings Protect components from damage

Enable proper flight characteristics Enable mounting of tail Enable control of flight surfaces Enable proper location of H&V tails Minimize detrimental effect of fuselage on wing lift Minimize drag

Provide volume for components Allow Proper location of C.G. Allow adjustment of C.G. location Allow extra volume for future component expansion

Structures Analysis: Wing Structure

12% max thickness wing Total max thickness is 0.9608” (XFOIL

determined) Highly cambered Planned construction Skinned hollow structure also allows

use of internal space Flaperon to be molded as part of wing Wing endplates manufactured

separately to allow internal wing access with removal

Structures Analysis: Fuselage Construction Requirements

Separate composite structure from wing

Supports tail Provide motor mount Light and durable Contain components

Propulsion System Selection

Propulsion System Selection

Create sufficient thrust to enable flight envelope Speed range 0 to ~50mph Thrust : Weight greater than 1:1 at 1kg weight level. (allowing future

platform expansion) >4min battery life at full throttle Reserve battery capacity to support future enhancements of

autonomous capabilities Minimal system weight Minimal system cost Minimum Electrical Noise System Safety

Propulsion System Selection

Determine required propeller characteristics Pitch picked based on max speed desired Propeller diameter picked to maximize efficiency also considering flow over the

fuselage/wing Motor selected to match propeller

Sized as small as possible to minimize weight Motor constant selected to match prop pitch to Vmax

Motor analysis done with concern for motor heating

Motor controller chosen for motor Amax

Battery selected to match system Series cells selected to create >=12V system Battery selected to match single pack to amp/hour requirements Manufacturer/design selected to allow required discharge rate with safety

margin to prevent battery damage

Propulsion System Selection

Single motor Small Brushless DC motor. Peak Watts 120 peak current 11amps Light, closely matches required motor constant (RPM/Volt), high quality components, small frame,

integrated gearbox Inefficient (~58%-67%) in operational envelope. Expensive, requires mount design

Single Prop 12” diameter 8” pitch Large enough to provide flow over wing, flow around fuselage, and flow over control surfaces at low

flight speeds. May be user safety risk, may be reliability/durability issue during landings

Speed Controller 20 Amp max speed controller (12g weight)

Battery 800 mAh/cell. 3.7 V/cell 20 C max discharge rate Provides 11.1 V and 4800 mAh total capacity Provides >8min run at full throttle/max load (vertical acceleration). Provides ~3 hours life at level

cruise (~33% throttle position at Clopt) Easily expandable for future research

Airfoil Selection

Airfoil Selection: Considerations

Design Considerations Maximize lift Maximize airfoil efficiency (Cl/Cd) Maximize stability Maximize flight envelope

AOA Velocity range

Initial metrics Flight velocity: V=30 mph=44 ft/s Lower limit on flight velocity: V=15 mph=44 ft/s Angle of attack: α=5° Elevation/atmospheric conditions: sea level @ standard atmospheric conditions

Density: =0.00237 (slug/ft3)

Kinematic Viscosity: =3.62x10-7 (lbf*s/ft2) Ambient temperature: Ta=528.67 °R Specific heat ratio: =1.4 Gas constant: R=1716 (ft*lbf/slug*°R)

Chord length: c=6 in=0.5 ft Wing span: b=20 in=1.667 ft Planform area: S=0.833 ft2

Required lift: L=9.81 N=2.205 lbf

Airfoil Selection: Initial Selection Process/Criteria

Initial metrics used to calculate non-dimensionalized operating conditions Equations

NASG and UIUC airfoil databases searched for any low Reynolds number airfoil listings

XFOIL viscous analysis used to analyze each of 160 possible airfoils discovered in the databases using the operating conditions relating to initial metrics Re=144033, M=.039, a=5° Record airfoil parameters

3 airfoils selected for secondary analysis based on their high Cl and Cl/Cd values Selig S1223 (Best Cl) Eppler E62 (Best efficiency) Selig S1210 (Balance of Cl and Cl/Cd)

Vc

Re aRT

VM

Airfoil Selection: Secondary Analysis

XFOIL utilized to obtain Cl data for the highest lift airfoil (S1223) over a range of AOAs from 0° to just beyond Clmax for 1 mph increments between V=15 mph and V=30 mph

Apply the following equations to find the maximum lift at each velocity (correcting for finite wing)

The maximum lift is found to be too low for our application Second iteration of metrics required to increase lift

c=8 in=0.667 ft Effective span (tip-to-tip span)-(4in): b=25.5 in=2.125 ft

Reynolds number recalculated for new metrics and XFOIL utilized to obtain performance data (Cl, Cd, and Cm) on the three airfoils

Performance data analyzed with respect to lift production and efficiency

Re1

A

CC

Cl

lL

180LL CC cbVCL L

2

2

1

Airfoil Selection: Results

Maximum Corrected Lift Comparison (Assuming a 20% Increase in Lift Due to Winglets)

0

100

200

300

400

500

600

700

800

900

1000

1100

1200

1300

1400

15 17 19 21 23 25 27 29

Velocity (mph)

Lif

t (g

) Selig S1223

Eppler E62

Selig S1210

Airfoil Selection: Results Continued

Re 192044 (V=30 mph)

0

0.5

1

1.5

2

2.5

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09

Cd

Cl

Selig S1223 Eppler E62 Selig S1210

Re 192044 (V=30 mph)

0

0.5

1

1.5

2

2.5

-10 -5 0 5 10 15

(deg)

Cl

Selig S1223 Eppler E62 Selig S1210

Airfoil Selection: Results Continued

Efficiency vs AOA Study Re=192044

0

20

40

60

80

100

120

-10 -5 0 5 10 15

(deg)

Cl/C

d

Selig S1223

Eppler E62

Selig S1210

Airfoil Selection: Corrected Results

Corrected Lift Comparison At Most Efficient AOA for Each Velocity(Assuming a 20% Increase in Lift Due to Winglets)

0

100

200

300

400

500

600

700

800

900

15 17 19 21 23 25 27 29

Velocity (mph)

Lif

t (g

) Selig S1223

Eppler E62

Selig S1210

Airfoil Selection: Final Decisions

Eppler E62 Best efficiency between of about 2° and 5°; however

efficiency drops precipitously beyond these values Very peaky and irregular efficiency, lift, and drag curves.

Narrow range of useable values: will decrease stability and make the plane more difficult to fly controllably Eliminated

Selig S1210 Highest corrected lift when operating at most efficient Moderate corrected Lmax in most likely flight envelope (V>=20

mph) Best efficiency for extreme values, moderate efficiency for

moderate values Smooth and regular efficiency, lift, and drag curves and wide

range of useable values (easier to fly) Selig S1223

Highest Lmax in most likely flight envelope Worst efficiency at moderate values, moderate efficiency at

extreme values Smooth and regular efficiency, lift, and drag curves and widest

range of useable values (easier to fly) Lowest corrected lift when operating at most efficient

Selection : Selig S1210

3rd iteration to metrics (based on airfoil selection)• =6° to 7° (most efficient AOA)• Plane mass: m=500 g (based on desired velocity range)

MAV Control System

MAV Control System: Previous Year’s

transmitter receiver

motors

Servos

4 Maximum

cameratransmitterreceiver

PPM motor control

NTSC video

Manual controller

Laptop with TV tuner

user

MAV Control System: Advantages and Disadvantages

Advantages of previous system: Easy to construct Cheap

Disadvantages of previous system: Limited motor control Very little vehicle data Fully manual control

MAV Control System: Proposed System

transceiver transceiver

motors

Servos

6 or 9 Maximum

cameratransmitterreceiver

RF-232 communication

NTSC video

Laptop with TV tuner

user

Flight controller

Sensor suite

MAV Control System: Potential Flight Controllers

UNAV Picopilot UNAV 3500FW O-Navi Phoenix AX Crossbow MNAV

Communication Interface RS-232 RS-232 RS-232 PPM, RS-232Number of PWM channels 4 7 6 9

CPU 8bit, 2MIPS 16 bit, 40MIPS 16 bit, 2.1MIPS Not providedSensor data rate 3Hz 30Hz 1, 10, 20, 50, 75 or 100Hz Not provided

GPS Yes Yes Yes YesBarometers static static static and dynamic static and dynamic

Angular Accelerometers Yes Yes Yes YesLinear Accelerometers No Yes Yes YesPower Consumption 0.3W 0.6W 1.01W 0.8W

Size 2" x 1" x 0.9" 4" x 2" x 0.8" 3.47" X 1.58" X 0.75" 2.25" x 1.8" x 0.44"Weight 31 g 34 g 45.3 g 33 g

Company Support No No Yes NoPrice $650 $2,500 $900 $1,500

Design of Experiments

Design of Experiments

Differential pressure sensor used in conjunction with a pitot tube to determine aircraft velocity

One differential pressure sensor on each wing to determine pressure difference which will yield AOA experimentally

Accelerometers placed in nose and tail to determine pitch, roll and yaw of aircraft during flight

Design of Experiments

Test Matrix of each component will be developed by end of Senior Design I

Entire team will define input variables and desired responses

Components will be tested to ensure they are in working order, as well as sending accurate information

Experiments will be carried out by entire team Statistical verification techniques will confirm

experiments are accurate and valid Lead Engineer will be responsible for document

control

Indication of Progress…

Platform design decided upon – COMPLETED Engineering metrics/product specifications – COMPLETED List of components and materials has been compiled – COMPLETED 3D CAD model of plane created (XFOIL has been used to narrow down

spectrum of airfoils, CAD model being generated) Foam model built based on concept generations (small scale model built) Tests are designed so that the components for use on the MAV are accurately

measuring required parameters (being looked into, design of experiments being generated)

Components have been ordered and/or are in the possession of the team (some components are being donated)

Components within possession of team are in the process of being tested via methods designed

Foam plane is fully functional from standpoint of flight and glide testing (scale model has been built)

Q & A