Critical Design Review - colorado.edu Pisano’s UPI autopilot simulator •Stability derivatives calculated from PowerFLOW System Architecture Deployment Mechanics ... Critical Design

Critical Design ReviewLeah Crumbaker

Jason FarmerJames Gordon

Matt Lenda

Jeffrey MullenScott TatumTravis SchafhausenKristina Wang

4 December 2008 Critical Design Review

Project CustomerProf. Eric Frew

Project AdvisorsProf. Bill Emery

Prof. Kurt Maute

1

Briefing Overview & Content

Critical Design Review

PurposeThis presentation will showcase the MADS critical design.

2

System Architecture

• Objectives

• Requirements

• System Design

• System Analysis

• Prototyping

Design Elements

• Mechanical

• Electrical

• Software

Manufacturing & System Integration

Plan

• Machining Considerations

• Assembly Flow Diagram

Verification & Test Plan

• Test plan

• Verification Items

Project Management Plan

• Org Chart

• WBS

• Risks

• Schedule

• Budget

• Special Facilities & Resources

Critical Design Review3

System Architecture

Design Elements

Manufacturing & System Integration Plan



Overview of Objectives


The mission of MADS is to develop a mechanism for a RECUV vehicle that can store and deploy four small sub-vehicles on demand.

This will provide a test platform that will allow the customer to develop communication protocols and algorithms concerning the optimization of deployment and utilization of multiple aircraft.

SystemArchitecture

4

Overview of Requirements


• The PV is the SIG Rascal 110

• The system shall carry and deploy 4 SVs in flight

• At least 1 flight-capable SV shall be delivered

Project Requirements

• Sub-Vehicle

• Minimum 15 minute endurance

• Minimum speed of 5 m/s

• Shall be controlled by the CUPIC autopilot

• Deployed on-demand (between 50 and 100 meters AGL)

• DM Impact on PV

• Minimum 15 minute PV endurance

• PV remains stable with deployment system in any configuration

System Requirements

DefinitionsPV = Primary VehicleSV = Sub VehicleDM = Deployment Mechanism

SystemArchitecture

System Design


1. Primary Vehicle (PV)• On-board PIC controls the Deployment Mechanism (DM) through Command and Data Handling (CDH) • Pilot controls the control surfaces

2. Deployment Mechanism (DM)• Consists of mounting point for the SV and linear actuator for pin-movement• Attached to the PV with bracketing system

3. Sub-Vehicle (SV)• CUPIC autopilot controls the control surfaces and motor settings through CDH• Payload is supplied with its own power

SystemArchitecture

6

System Flow Diagram


Primary VehicleSub Vehicle

Propulsion

Power Supply

Control

Power Supply

Control

Payload

Mechanism

Power Supply

Deployment Mechanism

CDH

CDHTX/RX

Propulsion

Command and Data Handling Center

TX/RX

Control SignalsPower Signals

Legend

SystemArchitecture

7

System Mass Budget


PV75%

SV18%

DM7%

Full System

Total Allowable Mass: 12.67 kgTotal Estimated Mass: 10.61 kgMargin: 2.06 kg (16.26%)

Sub VehicleElectronics – 36%Airframe – 24%Payload – 21%Motor – 16%DM – 3%

Deployment MechanismMechanism – 40%Actuators – 40%Upper Mount – 10%Lower Mount – 10%

Primary VehicleAirframe – 69%Engine – 21%Electronics – 5%Fuel – 4%Wiring – 1%

SystemArchitecture

SV Analysis


SV minimum 15 minute endurance, speed 5 m/sFurther Discussed in Electrical Design Elements

9

Item Value

Endurance 25.2 min (Safety Factor of 1.68)

Trimmed Flight Speed 10.1 m/s (Safety Factor of 2.20)

The SV design satisfies the system-level requirement

• SV is the SuperFly• SV is powered by Thunder Power RC 2S 7.4V, 2.1A*hr battery.

http://www.thunderpowerrc.com

SystemArchitecture

http://www.superflyrccom

SV & DM Analysis


SVs must be controlled by CUPICFurther Discussed in Software Design Elements

10

The SVs can be controlled by the CUPIC.

• With chosen power configuration, CUPIC autopilot can be powered for 15 minute endurance• Auto-pilot controllability evaluated by Bill Pisano’s CUPIC autopilot simulator

• Stability derivatives calculated from PowerFLOW

SystemArchitecture

Deployment Mechanics• Linear Actuators

• Force required: 4.54 N• Selected Actuator: Firgelli PQ12f Miniature Linear ServoStroke: 20mmForce: 18 N, SF = 4

SVs deployed on-demand (between 50 and 100 meters AGL)Further Discussed in Mechanical/Electrical Design Elements

The SVs can be deployed on-demand.

• Using Xbee communication, ground station able to send mission profiles on demand

• Using Xbee communication, ground station able to send deployment commands on demand

System Analysis


PV minimum endurance of 15 minutes with deployment systemFurther Discussed in Mechanical Design Elements

11

The PV will have an endurance of 15 minutes.

• Gasoline engine• Zenoah G26 EI

• Using flight data from CFD• Stock 15 oz tank = 39 min* endurance

* At 6000’ Altitude

SystemArchitecture

PV Remains Stable With Deployment SystemFurther Discussed in Mechanical Design Elements

• Linearized stability derivatives calculated with data from PowerFLOW• System is naturally stable except for spiral divergence

The PV remains stable with the deployment system.

Requests For Action


• Discussed in Mechanical Design Elements.

• Vibration Analysis: Simulate wind-initiated vibrations and turbulence, or cyclic vibrations to analyze the response of the Deployment Mechanism

• Requested by: Dr. Jean Koster• RFA Fulfilled: Performed a wide-range frequency response drop test & simulated aerodynamic loading drop test.

• Deployment Mechanism Prototype: Prototype and test the release mechanism to mitigate the risk of PV strike upon deployment.

• Requested by: Matt Rhode• RFA Fulfilled: A functional DM was prototyped and tested

SystemArchitecture


System Architecture

Design Elements




Mechanical Design Elements (MDE)

Design-To Specifications


•The SVs shall be deployed on demand

•The DM shall weigh no more than 13 g

•The DM shall be mounted on a rod (the bracket) capable of withstanding the expected loads.


•The DM shall not move the center of gravity of the System past [43 cm, 0 cm, -5 cm] with a margin of +/- 2 cm wrt the nose of the PV

•The mounting bracket shall be removable to minimize transport dimensions

Mounting Locations

• The deployment system shall not decrease the stability and handling characteristics of the PV such that it cannot be flown by an experienced RC pilot during all mission phases.

• The PV shall have a minimum flight endurance of 15 minutes with the deployment system.

Propulsion/Aerodynamics

MDE

14



Design


Deployment Mechanism Design• Actuator pulls a pin

• Pin removes attachment to SV• DM weighs 9g• DM mounted to an aluminum beam

Deployment MechanismMDE

• The SVs shall be deployed on demand.

• The DM shall weigh no more than 13 g

• The DM shall be mounted on a rod (the bracket) capable of withstanding the expected loads.

Deployment Mechanism Design-To Specs

Prototype & Testing Results• Under vibrations from 0 Hz to 150 Hz, successful deployment 121/124 trials

• Confidence of 95 % in vibrations• During simulated aerodynamic loading, successful deployment 20/20 trials

• Confidence of 99 % in aerodynamic loading

16

Design Path



17

Force Analysis



• Simulated “Worst Case” loads derived from aerodynamics of the sub-vehicle

• Lift = 12 N, Drag = 12 N• Friction (Al on Al) μs = 1.15

• Estimated friction force Fs = 19.44 N• Measured Force: Freq = 18 N• Decreased friction by filing and greasing the pin

• μs = 0.3• Freq = 4.54 N

Load

DM

Actuator Choice: Firgelli PQ12f Miniature Linear ServoStroke: 20mmForce: 18 N, SF=4Pull time: 3 seconds

• Simple linear motion• Contains potentiometer to verify pin is pulled

http://www.firgelli.com/products.php

18

Design Evolution


PDR Design• Difficult to manufacture• Large Heavy mass• SV can roll or yaw once the pin is pulled• Pin has possibility of jamming in the DM

PDR

CDR Design• Easy to manufacture (4 hours per DM)• Smaller SV Bracket Less added mass• Constrains roll and yaw once the pin is pulled• Minimizes possibility of jamming

CDR


19

Requests For Action


• Vibration Analysis: Simulate wind-initiated vibrations and turbulence, or cyclic vibrations to analyze the response of the Deployment Mechanism

• Requested by: Dr. Jean Koster• RFA Fulfilled: Performed a wide-range frequency response drop test & simulated aerodynamic loading drop test.

• Deployment Mechanism Prototype: Prototype and test the release mechanism to mitigate the risk of PV strike upon deployment.

• Requested by: Matt Rhode• RFA Fulfilled: A functional DM was prototyped and tested

20


101

102

103

5

10

15

Frequency [rad/s]

Fo

rce

[N

]

Maximum Force to Pull the Pin

Experimental Data

Best Fit Line

Vibration Test



21

Testing Results• Vibrations from 0Hz to 150Hz

• Vertical vibrations chosen due to most common mode in flight• Deployed 121/124 times.

• 95% confidence for non-aerodynamic loads in a vibration environment • Failed at 5Hz

• Theory: Large amplitudes of vibration caused the hinge joint between actuator and pin to vibrate beyond the tolerances of the DM• Pin attachment re-designed from a hinge to a solid joint

• Provides a SF 3.46 to 1.75 on the actuator pull force

51.50044.0 ffF

Aerodynamic Load Test



22

Drive Test• Tested DM functionality at flight conditions in order to verify that it could function in such an environment

• Speed: 15m/s

Data Collected:• Beam strain and twist angle• Beam tip acceleration

Conclusions• 20/20 Successful deployments under frequencies of 15 Hz• Observed post deployment SV motion• SV oscillates between AoA of -2 and 9 degrees

• Did not affect success of deployment

0 1 2 3 4 5-4

-2

0

2

4

6

8

10

Time [s]

An

gle

[d

eg

]

Angles as a Function of Time

Twist Angle

SV AoA

SV Released


Mounting Locations

Mounting Location Design


• The DM shall not move the center of gravity of the System past [43cm, 0cm, -5cm] with a margin of +/-2cm wrt the nose of the PV

• The mounting bracket shall be removable to minimize transport dimensions

Mounting Location Design-To Specs

MDEMountingLocations

Mounting Design1. Aluminum bars for SV mounting2. Aluminum reinforcement inside the PV fuselage3. Removable design4. Shifts the cg to [43.5cm, 0cm, -5.833cm]

24

Mounting Design Path



25

• CFD analysis of the SV performed to obtain worst-case loading scenario at 25m/s and 10°• Safety factor of 1.5 was added onto worst-case scenario

Mounting locations determination:1. Numerical Integration Drop Model2. Strongest locations on the PV fuselage

More analysis in appendix

ANSYS Model



Material Empirical Yield ANSYS FEA Results

Plywood 45 MPa 45 MPa

Balsa Wood 10.8 MPa 3.3 MPa

Aluminum 280 MPa 279 MPa

26


Propulsion and Aerodynamics

Design


Propulsion &AerodynamicsMDE

• The system shall have a minimum flight endurance of 15 minutes.

• The deployment system shall not decrease the stability and handling characteristics of the PV such that it cannot be flown by an experienced RC pilot during all mission phases.

Propulsion & Aerodynamics Design-To Specs

• PowerFLOW model of aerodynamics and stability

• System is longitudinally and laterally stable

• Entire system with SVs mounted• Vstall = 8.2 m/s• Takeoff Ground Run = 14.8 m

• Gasoline engine• Zenoah G26 EI

• Propeller• APC 16x8

• Using CFD performance• 15 oz tank = 39 min (SF = 2.6)

* At 6000’ Altitude

28

http://www.towerhobbies.com

http://www.modelflight.com.au

Design Path



29

System CFD Model



-10 -8 -6 -4 -2 0 2 4 6 8 10-0.2

0

0.2

0.4

0.6

System Aerodynamic Data at Re = 680940

Angle of Attack [deg]

Co

effic

ien

t

Lift

Drag

Moment

Stall

0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1 0.11-0.2

0

0.2

0.4

0.6

System Drag Polar at Re = 680940

CD

CL

• Key Aerodynamics• Stall AoA: 10°• L/Dmax: 7.63

30

System Stability Model



31

-25 -20 -15 -10 -5 0 5-20

-10

0

10

20

Poles of Longitudinal Modes

Real Axis

Ima

gin

ary

Axis

Short-Period

Phugoid

-0.7 -0.6 -0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 0.2-4

-2

0

2

4

Poles of Lateral Modes

Real Axis

Ima

gin

ary

Axis

Roll Subsidence

Spiral Divergence

Dutch Roll

Mode Value

Short Period 4.45 Hz

Phugoid 0.015 Hz

Dutch Roll 0.53 Hz

Roll Subsidence 0.10 Hz

Thrust Model & Flight Performance



32

PV Propulsion System Specifications

Propeller APC 16X8

Propeller Efficiency 0.7

Engine HP (Sea level) 2.7 hp

Engine HP (Boulder) 1.95 hp

Propeller RPM (Boulder) 8000 rpm

Power Available 1.37 hp

Pitch Speed 60 mph

Static Thrust 13 lb1

Thrust-to-Weight 0.6

1. http://peakeff.com/CalcTP.aspx2. http://h1069809.hobbyshopnow.com/ProdInfo/Files/ZEN_Manual_%20New-03-02-

2007.pdf

• Endurance• PowerFLOW Aerodynamics• Fuel Consumption 0.5 oz/min2

• Stock 15 oz tank• Breguet’s Equations• 39 min endurance• SF = 2.6

• Flight Envelope• PowerFLOW Aerodynamics• Thrust Model• System Mass• Vstall = 8.2 m/s• Takeoff Ground Run = 14.8 m


System Architecture

Design Elements




Electrical Design Elements (EDE)



• The command for the SVs to deploy shall be given by the GS when deployment is desired.

• The command for each SV release shall be a wireless signal sent from the GS to the CDH unit on the PV.


• The SV shall be capable of taking GPS data.

• The SV shall have a power supply of 7.4V or 11.1V with no more than 10A current draw.

Sub-Vehicle

• The PV shall have a minimum endurance of 15 minutes

• The PV shall be controlled by an experienced RC pilot

Primary Vehicle

EDE

34


Deployment MechanismElectrical Design

Design


On-demand deployment requires two elements for the DM electrical design:1. Wireless communication for commanding and telemetry: CUPIC hardware

• CUPIC hardware chosen because custom board would need Xbee and GPS2. Actuator and control circuit to pull the pin and deploy the SVs: Firgelli PQ12f Linear actuator and

custom controller utilizing the ADG888 DPDT Switch

Image Courtesy: Bill Pisano http://www.firgelli.com/products.php

http://www.analog.com/static/imported-files/data_sheets/ADG888.pdf

• The command for the SVs to deploy shall be given by the GS when deployment is desired.

• The command for each SV release shall be a wireless signal sent from the GS to the CDH unit on the PV.

Deployment Mechanism Electrical Design-To Specs

36

DM Electrical DesignEDE

DM Design Path



37

Actuator Controller Design


• Diodes provide back-EMF protection• Adjacent switches are shorted to ensure proper current flow• Once wireless command is received by the PIC on the PV, the PIC sends the logic 0 (0V) or 1 (+5V) command to the chip to change the state of the DPDT switches

Required Hardware:1. PQ12f Actuator (4)2. ADG888 DPDT IC (2)3. 1N345A Diode (8)

Controller Schematic

Controller Board Layout

38



Sub-VehicleElectrical Design

Design


•The SV shall be capable of taking GPS data.

•The SV shall have a power supply of 7.4V or 11.1V with no more than 10A current draw.

Sub-Vehicle Electrical Design-To Specs

Primary SV Design Element: • Control and Propulsion Units’ Power Supply

• Thunder Power RC 7.4V, 2.1A*hr 2-Cell Li-Po

Secondary SV Design Elements:• Elevon Servos

• Blue Bird BMS-380 4.8V, 57oz-in Metal

• Electronic Speed Controller• Thunderbird-18 18A, 15V maximum

• Payload• Supplies own power

SV Electrical DesignEDE

*Interface Legend In Appendix

SV Power Design Path


EDESV Electrical

Design

41

SV Thrust Model


• The SV propulsion and control units’ power supply is a strong function of the motor power draw.• Required thrust obtained from an free body diagram.

EDESV Electrical

Design

-10 -5 0 5 10 15 20 250

2

4

6

8

10

12

[deg]

TR

[N

]

Sub-Vehicle Required Thrust from CFD Model

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

Superfly Max

0 1 2 3 4 5 6 7 8 9 100

1

2

3

4

Sub-Vehicle Original Linear Thrust Model (VB = 7.4,

total = 60%)

IM

[Amp]

T [N

]

V

= 6 m/s

V

= 8 m/s

V

= 10 m/s

V

= 12 m/s

V

= 14 m/s

42

-2 0 2 4 6 8-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

Sub-Vehicle Forward Force and Motor Current (VB = 7.4V)

Motor Current [A]

Fo

rce

[N

]

V

= <5 m/s

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

-2 0 2 4 6 8-0.5

0

0.5

1

1.5

2

Sub-Vehicle Empirical Thrust (VB = 7.4)

Motor Current [A]

T [N

]

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

• A dynamic thrust test was performed in the ITLL wind tunnel in collaboration with Prof. Dale Lawrence

• Speeds from 0 to 20 m/s• AoA = 6° (L/Dmax for the SV)• Percent differences from 2 – 5% for thrust model based on CFD• Test validated CFD model!

SV Flight Configuration


EDE

Item Value

Battery 7.4V, 2.1A*hr

Required Motor Current 2.99A†

Maximum Allowable Motor Current 4.98A†

SV Mass w/o Payload 388g

Angle of Attack 6° [ ~(L/D)max ]

Allowable Payload 87g†

Achievable Endurance w/Payload 25.2min†

Trim Flight Speed w/Payload 10.1m/s

Stall Speed w/Payload 6.07m/s

SV Electrical Design

†Verified by flight test

• The thrust model is adjusted to fit the wind tunnel test data

• Design configuration gives:• Endurance Safety Factor: 1.68• Flight Speed Safety Factor: 2.20

43


Primary VehicleElectrical Design

Design


• The PV shall have a minimum endurance of 15 minutes

• The PV shall be controlled by an experienced RC pilot

Primary-Vehicle Electrical Design-To Specs

EDEPV Electrical

Design

Two independent subsystems on the PV:1. Propulsion and control

surface management.2. Engine ignition.

All parts were selected with RECUV pilot to ensure compatibility and functionality.

1.5A*hr batteries provide sufficient power for >15 minutes system flight.

*Interface Legend In Appendix


System Architecture

Design Elements




Software Design Elements (SDE)



• The GS shall send the SV release command to the PV

• The GS shall communicate with the aircraft using the XBee GS

• The GS shall have real time GPS processing for 5 CUPICs

Ground Station Software

• The CUPIC on the SV shall be capable of implementing a given deployment routine

• The CUPIC on the SV shall stream telemetry data to GS at 10 Hz

• The CUPIC on the SV shall control two elevons

• The CUPIC on the SV shall control the throttle

Sub-Vehicle CUPIC Software

• The CUPIC on the PV shall take release commands which must control 4 actuators for deployment

• The CUPIC on the PV shall confirm that each actuator has moved after commanded to do so

• The CUPIC on the PV shall stream GPS data back to the GS at 10 Hz

• The CUPIC on the PV shall stream potentiometer data back to the GS at 10 Hz

Deployment PIC Software

SDE

47

Communication and Data Flowchart


SDE

Primary Vehicle

CUPIC

XBee Radio

Actuators

Realterm &ActiveX

Ground Station GUI (MATLAB)

Ground Station

XBee Radio

Potentiometers

Wireless CommunicationWired Communication

Legend

Sub Vehicle

GPS

CUPIC

XBee Radio

Servos/Throttle

GPS

48


Ground Station Software Design

Design


Current GS Design:• Communicates with and parses data from up to 5 nodes

Added GS Functionality:• Add capability to communicate with PV PIC to deploy SVs on demand• Add capability to communicate with SV CUPIC to initiate deployment sequence• Receive potentiometer data

New Functions:• function deploy(actuator_number, src, evt)

• Initiates deployment sequence for a given sub-vehicle and actuator

• function checkDeployment(potentiometer_voltage)• Evaluates potentiometer data to confirm pin has pulled

Ground Station Software DesignSDE

• The GS shall send the SV release command to the PV

• The GS shall communicate with the aircraft using the XBee GS

• The GS shall have real time GPS processing for 5 CUPICs

Ground Station

50


Sub-Vehicle Software Design

Sub Vehicle Design


SDE Sub-VehicleSoftware Design

•The CUPIC on the SV shall be capable of implementing a given deployment routine

•The CUPIC on the SV shall stream telemetry data to GS at 10 Hz

•The CUPIC on the SV shall control two elevons

•The CUPIC on the SV shall control the throttle

Sub-Vehicle CUPIC Design-To Specs

Current CUPIC Design:• Streams telemetry data to GS• Controls surfaces/throttle

Added CUPIC Autopilot Functionality:• New loop created before deployment using interrupt system to receive commands• Preset elevon and throttle control sequence initiated upon parsing of deployment command from ground station• Upon completion of deployment loop, normal autopilot loop starts

52


Deployment PIC Software Design

Deployment PIC Design


Deployment PIC Software DesignSDE

• The CUPIC on the PV shall take release commands which must control 4 actuators for deployment

• The CUPIC on the PV shall confirm that each actuator has moved after commanded to do so

• The CUPIC on the PV shall stream GPS data back to the GS at 10 Hz

• The CUPIC on the PV shall stream potentiometer data back to the GS at 10 Hz

Deployment PIC Software Design-To Specs

Modified Deployment PIC Functionality:• CUPIC hardware used, but autopilot software removed –still uses interrupt system to receive commands• Preset actuator control sequence initiated upon parsing of actuator deployment command from ground station• Potentiometer data for each actuator as well as GPS data is streamed to ground station

54

Fault Protection


Failed Deployment• Determined by actuator potentiometer data• Deployment routine is terminated

Sub-Vehicle Range• Landing mode is activated on autopilot (code already exists)

Corrupt Command (stretch goal)• Command confirmation system

• Ground Station sends command once per second• When command is received, it is transmitted back for validation• If command is incorrect, command is resent

55

SDE


System Architecture

Design Elements




Manufacturing Considerations

57Critical Design Review

Machined Parts: Deployment Mechanism

Machine Number of Parts Time estimate (hours)

Horizontal Bandsaw 2 1

CNC Mill 4 6

Drill Press 3 1

Total (x4 DMs) 36 parts 32 hours

Purchased Parts

Sub-System Number of Parts MADS Cost

Primary Vehicle (2) 55 $400.00

Sub-Vehicle 10 $691.50

Deployment Mechanism 15 $334.50

Ground Station 4 $140.00

MSIP

Assembly Flow Diagram



• DM body (top)• Sleeves (2)• Screws (2)• Pin

DM (bottom) DM strut

Actuator

Hot glue and screw to SVcanopy.

Screw DM onto strut. Screw strut through fuselage of PV & reinforcements. Held in with

set screws.

U-bolt to strut.

Full System• PV, SV, DM

Ground Station

• Laptop• RC Transmitter• Xbee Pro GS

Communication

Sub-vehicle• Battery• CUPIC• Motor• Propeller• Servo (2)• Z bent connector

Primary Vehicle

• CUPIC• Engine• Fuel• Internal aluminum reinforcements (2)

DM top and bottom held

together by pin.

Pin attaches to actuator with lock washer and screw.

MSIP


System Architecture

Design Elements




Verification ItemsVerification &

Test Plan


Primary Vehicle

SIG Rascal 110 Deploy Altitude (50-100m) V = 15m/s at deployment E = 15minMax Payload = 4.71kgMax Mass = 12.67kg Controlled by RC Pilot Stability

Sub-Vehicle

Flight CapableMin V = 5m/s E = 15minMax Mass = 485gMax Payload = 0.1 kg Stability


Carry 4 SVs for 15 minutes Deploy SV on Demand within 3 secondsMax Mass = 0.013kg

CDH

Sample GPS data at 10 Hz Stream data at 10 Hz Control SV elevons Control SV throttle Stream potentiometer data from PV at 10 Hz Interpret data

SystemFull-System Fit and Functionality

PV – DM interface The system shall not alter from [43 cm, 0 cm, -5 cm] more than of +/- 2 cm in the X, or Z direction DM in flight shall not exceed the material limits of the PV structureMount within specified area

DM – SV interface DM shall not decrease the structural integrity of the SV

Flight environment Operate in the temperature range of 0-40 degrees Celsius. Shall not operate if winds exceeds 7 m/s

Velocity & altitude at

drop

Testing Overview


Individual Component

Tests

Interface Testing

between Sub-systems

Full System Integration

& Static Test

Flight Testing &

Verification

PV SV DM CDH

EnduranceBattery

Test

Static Engine

Test

EnduranceBattery

TestSV Mass

Static Bench Test

Car Test: DM

(pull pin)

CU PIC Ground

TestGS Test

Car Test: DM

(Actuator)

PV-SV Mounting

PV-DM Modulus

PV-DM Structural Integrity

Static Full System Test

Total Mass

Full System Flight Test & Operating Environment Measurements

PV Stability

SV Stability

Waypoint & Loiter

DM functionality

CDHfunctionality


Test Plan


Test Measurements EquipmentModel

VerifiedReq.

Verified

Endurance • Battery capacity• Time to deplete battery w/ constant WC power draw

• Battery• Electrical comps.• Resistor

Endurancemodel

0.SYS1

Verify functionality of CUPIC and the GS

Collect GPS Data:• Transmission rate• Receiving rate• Compute velocity, acceleration,distance

• CUPIC• Xbee GS• Software

NA 0.SYS4

Structural Integrity of PV & DM Interface Un/Modified

Verify ANSYS model:• Forces applied• Moment applied• Location of forces

• Weights (WC load)• PV• DM• Strut reinforcement

ANSYS model 0.SYS5

DM Release on Demand (Car Test)

• 2D accelerations• Stress• Strain• Torsion• High-speed video

• Strut• SV (w/ weights)• DM• Sensors• Pickup truck• Ladder

Drop model 0.PRJ1

Verify w/Flight Test:• System stability• Full system functionality

• PWM duty cycles• Roll/Pitch/Yaw rates• X/Y/Z acceleration rates• GPS data

• 6DOF IMU• PV, SV, DM, GS assembled & integrated• Oscilloscope• Pilot (2)

CFD & stabilitymodel

0.SYS10.SYS20.SYS50.SYS6



System Architecture

Design Elements




PMP: Organizational Chart


MADS

Travis SchafhausenProject Manager

Leah CrumbakerSafety Engineer Scott Tatum

Chief Financial OfficerKristina WangWebmaster

Leah CrumbakerSystems Engineer

AerodynamicsJeff Mullen

Kristina WangLeah

Crumbaker

Avionics/ ElectronicsMatthew

LendaJeff Mullen

Travis Schafhausen

SV DesignKristina Wang

Matthew LendaTravis

Schafhausen

Mechanics/ DeploymentScott Tatum

Jason FarmerLeah

Crumbaker

Software James Gordon

Matthew Lenda

Jeff Mullen

FabricationJason Farmer

TestJeff Mullen

Travis Schafhausen

Structures Jason FarmerScott Tatum

James Gordon

Eric FrewCustomer

Kurt MauteWilliam EmeryPAB Advisors

64

PMP

Work Breakdown Structure


WBS Manufacture

System SV DM CDH

Assembly

Structural Modification

DM Integration

Assembly Mechanism Machining

Assembly

DM Deployment Microcontroller Programming

Deployment Command and Confirmation

SV CUPIC Deployment Routine

PMP

Work Breakdown Structure


System SV DM CDH

Static Structural Load Test

Static Thrust Test

Propulsion System

Endurance

Power System Endurance

Control System Endurance

Bench Flight

Pilot Familiarization

Taxi Test

Half Configuration

Full Configuration

Endurance Test

PMP

WBS Test

Power System Endurance Test

RC Control Functionality

CUPIC Control Functionality

Control Switching

(RC/CUPIC)

Endurance Test

Flight Velocity Verification

Autopilot Control Test

Static Deployment Test

Dynamic Deployment 1

(Drive Test)

Vibration Test

Dynamic Deployment 2

(Drive Test w/Actuator)

Deployment Confirmation

Deployment Confirmation

In-Flight Deployment

Bench Flight Bench Flight CUPIC Sensor Functionality

XBee Communication (CUPIC & GS)

GS Multi-Node Capabilities

Deployment Command

Risks


PV Unstable or Uncontrollable with

DS

•Full system is stable

•DM design mitigates partial or accidental deployment resulting in uncontrollable configurations

•Standard configurations controllable

SV Strikes PV on Deployment

•Drop Model predicts clean deployment

•Drop Test demonstrates clean deployment

•Known SV deployment state with CUPIC

Mass Change Alters Stability or Endurance

•Mass budget set

•Full system is stable

•Large mass margin

DM Failure

•Fault Tolerance –default state closed, unpowered until deployment

•DM fully constrained – no partial deployment

•Vibration test

PMP

Flutter

•Cannot be modeled or predicted

•Certain modes detected during vibration test might be catastrophic during flight

PV Structural Integrity

• Assumptions in the boundary conditions of the FEA model for the structure may not be valid

• Non-destructive static load test will be performed

67

Schedule


PMP

68

Fiscal Budget


Cost Estimates

Product Cost ($) # Total ($) Notes

Primary Vehicle

Aircraft 0.00 1 0.00

Provided by RECUV

Servos/Electronics 0.00 1 0.00

Engines 0.00 1 0.00

Fuel System 0.00 1 0.00

Field Equipment 0.00 1 0.00

CUPIC 400.00 1 400.00

Subtotal $ 400.00

Sub Vehicle

Airframe 50.00 4 200.00

Motor 16.50 1 16.50

Servo/Electonics 75.00 1 75.00

CUPIC 400.00 1 400.00

Subtotal $ 691.50


Mountings 54.50 1 54.50

Linear Actuator 70.00 4 280.00

Subtotal $ 334.50

Ground Station

RC Transmitter 0.00 1 0.00 Provided by RECUV

XBee Pro Ground Station 140.00 1 140.00

ICD2 Programmer 0.00 1 0.00 Available from ITLL

Laptop 0.00 1 0.00 Team Supplied

Subtotal $ 140.00

Margin (30%) $ 1,200.00

Total $ 2,766.00

Available $ 4,000.00

Balance/Additional Margin $ 1,234.00 30.9%

10%

17%

8%

4%

30%

31%

MADS System

Primary Vehicle Sub Vehicle

Deployment Mechanism Ground Station

PDR Margin (30%) Balance/Additional Margin

69

PMP

$1,234.00 additional margin available

Specialized Facilities & Resources


If required, the following facilities and resources may be utilized:

RC FieldPOC: Eric FrewRECUV Table Mountain Flight Range has an area for flight testing.

RECUV RC PilotPOC: Tom AuneTom has offered to pilot the system for flight testing pending availability.

Frank DilatushFrank has provided an abundance of useful information regarding RC planes and suggested our final SV choice (the SuperFly). He has also offered additional time and resources to our team during the critical design and testing phases of the project.

70

PMP

Conclusions


• The critical design has been shown to fulfill all project and system-level requirements

71

Acknowledgements


Prof. Bill Emery and Prof. Kurt Maute – Our faculty advisors have given useful feedback throughout the design process.Prof. Dale Lawrence – Dr. Lawrence supervised our wind tunnel testing and allowed us to use his equipment for data collection. Frank Dilatush – Frank has provided an abundance of useful information regarding RC planes and suggested our final SV choice. He has also offered additional time and resources to our team during the critical design and testing phases of the project.Eric Frew – Our customer has offered many hours in helping us develop our project goal and requirements.Georg Pingen – Georg assisted with CFD analysis by offering help with PowerFlow.David Halko– The designer and manufacturer of SuperFly RC planes has provided additional specifications and information about his products and offered future help with the project.Trudy Schwartz – Trudy has been a great help in our consideration of electromagnetic deployment during preliminary design. She has guided us in our calculations and component search. Bill Pisano – Bill has helped a great deal in aiding our understanding of the CUPIC.Tom Aune - Tom has helped with the selection of PV flight componentsPAB – The members of the PAB have provided constructive criticism, feedback, and guidance throughout project definition, development, and design.

72


Questions?


System Architecture

Design Elements





System Architecture

Design Elements




Project Level Requirements


Req. # Description Verification

0.PRJ1 The deployment mechanism shall be able to carry and deploy four SVs in flight. 3.TST13.TST23.TST33.TST43.TST53.TST6

0.PRJ2 The PV shall be the RC aircraft SIG Rascal 110. Inspection

0.PRJ3 MADS shall deliver a minimum of one flight-capable SV in spring 2009. 0.TST2

0.PRJ4 All units of measurement used throughout the project shall be metric. Inspection

0.PRJ5 The PV and SVs shall abide by all rules and regulations enforced by the Academy of Model Aeronautics. Inspection

1.PRJ1 The SV shall be deployed from an altitude between 50 and 100 meters above ground level (AGL) with respect to Boulder, Colorado.

3.TST6

SystemArchitecture

System Level Requirements


Req. # Description Verification

0.SYS1 The PV shall have a minimum flight endurance of 15 minutes, +/- 1 minute, with the full deployment system attached.

1.TST11.TST21.TST3

0.SYS2 The SV shall have a minimum 15 minute, +/- 1 minute, post-deployment flight endurance. 2.TST22.TST6

0.SYS3 The SV shall have a minimum airspeed of 5 m/s, +/- 0.1 m/s. 2.TST5

0.SYS4 The SV shall utilize the GPS-integrated autopilot chips (CUPIC). 0.TST10.TST24.TST1

0.SYS5 The deployment system shall not decrease the stability and handling characteristics of the PV such that it cannot be flown by an experienced RC pilot during all mission phases.

0.TST40.TST5

0.SYS6 The deployment mechanism shall deploy each SV on demand. 3.TST1

0.SYS7 The PV shall land if an SV fails to release. Inspection

0.SYS8 The system shall only be flown in fair-weather conditions. 0.TST9

1.SYS1 The SV shall be capable of stable flight under autopilot control. 2.TST4

1.SYS2 The CUPIC shall sample data at TBD determined by January 15, 2008. 0.TST10.TST2

1.SYS3 The CUPIC shall stream data at a rate of TBD determined by January 15, 2008, to the ground station over a wireless network.

0.TST10.TST2

SystemArchitecture

System Performance


The system must fly for 15 minutesThe SV must fly for 15 minutes with

a minimum speed of 5 m/sThe SVs shall deployed on-demand

•Drag polar• Performance parameters• Stability

Aerodynamic Model

• Engine• Fuel tank size

Propulsion

•Drag polar• Performance parameters• Stability

Aerodynamic Model

• Thrust vs. Current

Thrust Model

• Compare model to reality

Dynamic Thrust Testing

• Blah

Risks

• Blah blahblah

Drop Model

• Drop testing• Vibration testing

Testing

78

SystemArchitecture

Detailed Mass Budgets


Mass Budget

Product Mass (kg) # Total (kg)Notes

Primary Vehicle

Airframe 5.500 1 5.500

Engine 1.520 1 1.520

Ignition Battery 0.120 1 0.120

Propeller 0.000

Fuel Tank 0.000

Fuel 0.320 1 0.320

Control Surface Servo 0.047 4 0.186

Throttle Servo 0.037 1 0.037

Receiver 0.010 1 0.010

Control System Battery 0.120 1 0.120

CUPIC 0.030 1 0.030

Wiring/Cables/Mountings/Etc 0.100 1 0.100Margin

Subtotal 7.943

Sub Vehicle

Airframe 0.110 4 0.440

Motor 0.048 4 0.193

Motor Controller 0.019 4 0.077

Propeller 0.006 4 0.022

Servo 0.016 8 0.125

Servo Connection Rod 0.006 8 0.048

Battery 0.095 4 0.380

CUPIC 0.030 4 0.120

DM Attachment 0.016 4 0.064

Payload 0.100 4 0.400

Subtotal 0.446 1.869


Upper Mounting Rod 0.195 1 0.195

Upper Mounting Reinforcement 0.026 2 0.052

Lower Mounting Rod 0.200 1 0.200

Lower Mounting Reinforcement 0.046 2 0.092

Lower Mounting Connector 0.020 1 0.020

Mechanism 0.040 4 0.160


Subtotal 0.795

Total 10.607

Available 12.670 30 min Endurance

Balance/Additional Margin 2.063 16.3%

79

SystemArchitecture

Subsystem Mass Budget Breakdown


10%10%

40%

40%


Upper Mounting Lower Mounting Mechanism Linear Actuator

69%

21%

4%

5%

1%

Primary Vehicle

Airframe Engine Fuel Servos/Electronics Miscellaneous

24%

16%

36%

3%

21%

Sub-Vehicle

Airframe Motor Servos/Electronics DM Attachment Payload

80

SystemArchitecture


System Architecture

Design Elements




Mechanical Design Elements (MDE)

Mechanical Drawing Tree


MDE

Mechanical Drawing Tree


MDE



SV Mounting Bracket


• Easier to Fabricate• Low weight size and design• Possible Masses:

•AL 2024 – 10.37 g•Mg Alloy – 6.29 g

•Dimensions:• L=1.25”• W=.5”• H=.625’

• Estimated total mass addition:•AL 2024 – 13 g•Mg Alloy – 8 g

• Secured through foam with screw, locking nut and custom washer to distribute load.

L

W

H

85


Material Trade Study


MaterialDensity

(lbs/in^3)Elastic

Modulus (ksi)Specific Modulus

(in)Compressive Yield

Strength(ksi)

Magnesium 0.067 6380 94813.49 21.176

Aluminum Alloy 0.102 10400 102362.20 33.214

High Density PE 0.035 120 3458.21 -

ABS Plastic 0.034 300 8875.74 -

Acrylic 0.042 230 5418.14 13.927

• Aluminum and Magnesium best choices • Aluminum chosen – highest specific Modulus

• SV can carry aluminum mounting bracket• Magnesium is too weak in compression• Readily available – easily machined

86


Deployment: Static Test


• Simulated “Worst Case” aerodynamic loads • Simulated “Worst Case” friction environment•Lift applied = 12 N, Drag applied = 12 N• Combined Force applied = 16.97 N = 3.8lb• Estimated friction (Al on Al) - μs =1.15•Estimated friction force -Fs =4.37 lbf

• Measured average force required to pull pin - Freq =4.0625 lbf• Can decrease friction by ¾ using oil - Freq =1.021 lbf = 16.336 oz• Torques required –

• With ½” servo moment arm Treq =32.6 in*oz• With 1” servo moment arm Treq =16.336 in*oz

• Chosen servo operates at 23 oz*in of torque, supplies a 1” wheel in kitCONCLUSION:

• After prototyping and testing, this DM will deploy on command

87


Drop Test #1 Set-Up


14cm

Y

X

V∞

22cm 34cm

Drag

PitchingMoment Z-Axis

Accelerometer

34cm

Y

Z

X-AxisAccelerometer

PitchingMoment

Lift

DM Mounting BracketStrain GaugeAccelerometerDM

Legend

Y-Z Plane

X-Y Plane

88


Vibration Analysis


Mode Value Deploy?

PV - Phugoid 0.13 Yes

PV - Dutch Roll 0.51 Yes

PV - Roll Subsidence 0.99 Yes

System – Short Period 4.45 Yes

System – Phugoid 0.015 Yes

System – Dutch Roll 0.53 Yes

System – Roll Subsidence 0.10 Yes

Beam Mode 1 6.662 Yes





Mode Value Deploy?

SV – Short Period 22.4 Yes

SV – Phugoid 0.074 Yes

SV – Dutch Roll 1.99 Yes

SV – Roll Subsidence 0.080 Yes

100

101

102

103

-30

-20

-10

0

10

Ma

gn

itu

de

[d

B]

Open Loop Bode Plot of Deployment System

100

101

102

103

-130

-120

-110

-100

Frequency [rad/s]

Ph

ase

[d

eg

ree

s]

89


Sub-Vehicle Drop Model


Methodology:• Force/Moment sum from PowerFLOW data• Numerical integration to calculate position, velocity, pitch angle, and flight path angle

Features:• Variable initial conditions• Deployment mechanism constraint in horizontal/vertical directions• Variable thrust/elevon input as a function of time• Collision detection

• Projections of collision areas made onto X-Z plane• Wing, wheel-casing, and non-deployed sub-vehicle boxed for efficiency• Deployed sub-vehicle uses simplified shadow as seen in relative motion plot• Collision detected by detecting intersecting line segments




-1.6 -1.4 -1.2 -1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4

-1.4

-1.2

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

Tail

Wing

Wing Strut

Wheel Strut

Wheel Casing

Upper SV

Relative Motion of SV to SIG

Horizontal Motion (m)

Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone

Initial Conditions:α ≈ α(L = W) ≈ 1.66o

V = 15m/stflight = 0.5stRelease = 0.083sδElevon = 0o

Thrust = 0 N

• Deployment was successful• No collisions detected




Initial Conditions:α ≈ α+Stall ≈ 22o


Thrust = 0 N


-1.6 -1.4 -1.2 -1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4

-1.6

-1.4

-1.2

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

Tail

Wing

Wing Strut

Wheel Strut

Wheel Casing

Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone




Initial Conditions:α ≈ α-Stall ≈ -20o


Thrust = 0 N


-2 -1.5 -1 -0.5 0 0.5

-2

-1.5

-1

-0.5

0Tail

WingWing Strut

Wheel StrutWheel Casing

Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone





V = 15m/stflight = 0.5stRelease = 0.036sδElevon = δElevon, +Max = 25o

Thrust = 0N

• Deployment was successful• No collisions detected• Vehicle stalls after 0.0560s

-1.2 -1 -0.8 -0.6 -0.4 -0.2 0

-0.6

-0.4

-0.2

0

0.2

0.4

Tail

Wing

Wing Strut

Wheel Strut

Wheel Casing

Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone





V = 15m/stflight = 3stRelease = 1.630sδElevon = 15o

Thrust = 0 N


-3.5 -3 -2.5 -2 -1.5 -1 -0.5 0 0.5 1

-3.5

-3

-2.5

-2

-1.5

-1

-0.5

0Tail

WingWing Strut


Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone





V = 15m/stflight = 0.5stRelease = 0.129 sδElevon = 5o

Thrust = 0 N


-1.2 -1 -0.8 -0.6 -0.4 -0.2 0

-0.6

-0.4

-0.2

0

0.2

0.4

Tail

Wing

Wing Strut

Wheel Strut

Wheel Casing

Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone





V = 15m/stflight = 0.5stRelease = 0.072sδElevon = -5o

Thrust = 0 N


-2 -1.5 -1 -0.5 0 0.5 1

-2.5

-2

-1.5

-1

-0.5

0Tail

WingWing Strut


Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone





V = 15m/stflight = 0.5stRelease = 0.072sδElevon = -15o

Thrust = 0 N


-3.5 -3 -2.5 -2 -1.5 -1 -0.5 0 0.5

-3

-2.5

-2

-1.5

-1

-0.5

0Tail

WingWing Strut


Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone





V = 15m/stflight = 0.5stRelease = 0.072sδElevon = δElevon, -Max = -25o

Thrust = 0 N

• No collisions detected• Vehicle stalled before deployment completed at 0.0440s

-1.2 -1 -0.8 -0.6 -0.4 -0.2 0

-0.6

-0.4

-0.2

0

0.2

0.4

Tail

Wing

Wing Strut

Wheel Strut

Wheel Casing

Upper SV



Vert

ical M

otion (

m)

Path

Start

End

Restricted Zone



Mounting Locations

Mounting Model FBD


• Option 2 was selected• More distributed load through both shear and normal stresses• 3 sections to analyze: Applied Forces, Reactions, and Stresses

V∞

σzz

τzxτzy

σyyτyx

τyz

σxx

τxz

τxy

SV Drag (Fdrag)

Lift (Flift)

Drag of Bracket (wb)

PitchingMoment(Mm) Vyz

Vzx

My

Reactions

YX

Z

Applied Forces From SV

Stresses

Modeling Issues:• This is an over determinate problem in order to resolvethe stresses at the attachment point• Finite Element Analysis (FEA) will be used to solve forstresses and ensure materials/reinforcements do notbreak in worst-case loading scenarios

101


Applied Loads


CFD analysis of the SV performed to obtain worst-case loading scenario at 25m/s and 10°

Item Worst-Case Value with SF = 1.5

Lift 50 N

Drag 10 N

Moment -3 Nm

-20 -15 -10 -5 0 5 10 15 20 25-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

SV Aerodynamic Data at Re = 272819


Co

effic

ien

t

Lift

Drag

Moment

Stall

102


Ideal Mounting Locations


-1.6 -1.4 -1.2 -1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4

-1.6

-1.4

-1.2

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

Tail

Wing

Wing Strut

Wheel Strut

Wheel Casing

Upper SV


Horizontal Motion (m)V

ert

ical M

otion (

m)

Path

Start

End

Restricted Zone

Mounting location determination:1. Numerical Integration Drop Model2. Strongest locations on the PV

fuselage

Results:• SVs naturally pitch away from the

PV upon deployment – confirmed in drive/drop test

• Upper beams mounted at 10°• Beams extend 43.5cm from

fuselage

103


Mounting Option Trade Study


Pros• No holes required to drill into PV

Cons• Difficult to reinforce existing structure without adding parasitic drag• All loads are shear

Pros• Loads are distributed through both shear and normal stresses• Easier to reinforce without affecting aerodynamics of PV• Vibration of beam will be damped out by structure of PV

Cons• Difficult to model• May interfere with control lines/electronics of PV

Pros• Loads are distributed through both shear and normal stresses• Beam takes majority of shear stress• Easy to reinforce without affecting aerodynamics of PV• Vibration of beam will be damped out by structure of PV

Cons• Difficult to model• May interfere with control lines/electronics of PV

1 2 3

104


Material Testing


• In order to determine flexural modulus (Eflexural) and the max bending stress (σmax) for the FEA model, 3-point bend test was performed

• 3-point bend tests are usually performed with brittle materials• Difficult to perform tension test with wood

• Slides out of grips• 2 specimens were tested two different with test mechanism parallel to perpendicular grains

• 1/8” plywood• 1/8” balsawood

• 4 samples were used for each test, results in a total of 16 tests

Samples after break test

Perpendicular Grain Balsa Wood Under Load

Parallel Grain Plywood Under Load

105


Material Testing


Material Orientation Eflexural σmax

Plywood Perpendicular 13.63 GPa 44.83 MPa

Plywood Parallel 4.77 GPa 22.19 MPa

Balsa Wood Perpendicular 3.07 GPa 10.87 MPa

Balsa Wood Parallel 107.28 MPa 931.09 kPa

0 1 2 3 4 50

500

1000

1500

Results for Plywood - Perpendicular Grain

Cross-Head Position [mm]

Ap

plie

d L

oa

d [N

]

0 1 2 3 4 5 60

200

400

600

Results for Plywood - Parallel Grain


Ap

plie

d L

oa

d [N

]

0 2 4 6 80

10

20

30

Results for Balsa Wood - Parallel Grain


Ap

plie

d L

oa

d [N

]

0 1 2 3 4 5 6 70

100

200

300

Results for Balsa Wood - Perpendicular Grain


Ap

plie

d L

oa

d [N

]

Experimental Data

Break Point

106


Structural Load Cases


FEM analysis 5 loading casesCase 1 Case 2

Case 3 Case 4

Case 5

107


Structural Stresses


Case 1 Case 2 Case 3 Case 4 Case 5

MaxPlywood Stress (MPa)

45 45 45 43 41

Max Balsa Stress (MPa) 3.3 3.3 2.8 3.3 2.8

Max Beam Stress (MPa) 279 279 279 278 278

108


Structural Forces Applied


• Loading scenario: Worst Case• V = 25 m/s•AoA = 10°• Safety Factor = 1.5

• Resulting Forces• Lift = 50 N• Drag = 10 N• Moment = -3 Nm

109



Propulsion and Aerodynamics

PV CFD


-10 -5 0 5 10 15-1

-0.5

0

0.5

1

1.5

PV Aerodynamic Data at Re = 408121


Co

effic

ien

t

Lift

Drag

Moment

Stall

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35-0.5

0

0.5

1

1.5

PV Drag Polar at Re = 408121

CD

CL

• Key Aerodynamics•Stall AoA: 12°• L/Dmax: 12.07

111


PV Stability


• Stable in all flight modes• Short-period is essentially non-existent •Phugoid—No oscillatory motion

•ω = 0.13 Hz, T = 7.70 s•Dutch Roll

•ω = 0.51 Hz, T = 1.97 s• Roll Subsidence

•ω = 0.99 Hz, T = 1.01 s• Spiral Mode

•ω = 0.05 Hz, T = 20.00 s

-90 -80 -70 -60 -50 -40 -30 -20 -10 0-0.4

-0.2

0

0.2

0.4


Real Axis

Imagin

ary

Axis

Short-Period

Phugoid

-7 -6 -5 -4 -3 -2 -1 0-4

-2

0

2

4


Real Axis

Imagin

ary

Axis

Roll Subsidence

Spiral Divergence

Dutch Roll

112


System Center of Gravity


y

x

z

Plane Mass (kg)CG

X-Location (cm)CG

Y-Location (cm)CG

Z-Location (cm)

Primary Vehicle 8.5 38.40 0 -3.96

Sub-Vehicle 1 .4 81.03 50 -7.2175

Sub-Vehicle 2 .4 81.03 -50 -7.2175

Sub-Vehicle 3 .4 61.03 50 -24.36

Sub-Vehicle 4 .4 61.03 -50 -24.36

mass

radiusmassCG

CG location of the System• X-Direction: 43.5691 cm• Y-Direction: 0 cm• Z-Direction: -5.8339 cm

113


System Moments of Inertia


Original MOI of the PV• Ixx = 2.6410 kg-m2 Iyy = 2.1045 kg-m2 Izz = 2.5985 kg-m2

...321 IIIIsystem 2mRII centerdisplaced

Addition of the SVs Changes MOI• Parallel Axis Theorem

• MOI of SV about the SV CG with component estimates and Parallel Axis Theorem• Find system MOI using original PV MOI and Parallel Axis Theorem

Updated MOI of the System• Ixx = 2.8366 kg-m2

• Iyy = 2.5166 kg-m2

• Izz = 2.6449 kg-m2

114



System Architecture

Design Elements




Electrical Design Elements (EDE)

Electrical Drawing Tree


EDE

116

EBD Wiring Key


EDE

Electrical Drawing: 1.0


EDE



EDE

Electrical Drawing: 1.1.1


EDE



EDE



EDE



EDE



EDE



EDE



EDE



EDE


Deployment MechanismElectrical Design

DM Force and Power Req’s


Item ValuePower Supply 5VDCCurrent Draw 250mA

Stroke/Pin Length 20mmRated Force (Speed) 15N (7mm/s)

Maximum Force (Speed) 18N (6mm/s)Maximum Speed 12mm/s

Operating Temperature -10°C to +50°C

• As chosen in MDE, the Firgelli PQ12f linear actuator will be used for the DM

• In order to control PQ12f, polarity reversal is needed• Firgelli recommends a DPDT switch that allows >250mA

Item ValueVDD to GND -0.3V to +6.0V

Allowable Peak Current 600mA (5V)Allowable Run Current 400mA (5V)

Impedance <0.6ΩOperating Temperature -40°C to +125°C

http://www.firgelli.com/products.php

http://www.analog.com/static/imported-files/data_sheets/ADG888.pdf

129


DM Level-2 EBD


EDEDM Electrical

Design


Sub-VehicleElectrical Design

SV Thrust Model


• The SV propulsion and control units’ power supply is a strong function of the motor power draw.

M

pmM

M

pmMM

IV

VIfT

VITVP

)(

-10 -5 0 5 10 15 20 250

2

4

6

8

10

12

[deg]

TR

[N

]


V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

Superfly Max

0 1 2 3 4 5 6 7 8 9 100

1

2

3

4


total = 60%)

IM

[Amp]

T [N

]

V

= 6 m/s

V

= 8 m/s

V

= 10 m/s

V

= 12 m/s

V

= 14 m/s

• Required thrust obtained from an free body diagram.

cos

DTR

132


Wind Tunnel vs. CFD


• A dynamic thrust test was performed in the ITLL wind tunnel in collaboration with Prof. Dale Lawrence• Speeds from 0 to 20 m/s• AoA = 6° (L/Dmax for the SV)• Percent differences from 2 – 5% for thrust model based on CFD• Test validated CFD model!

-2 0 2 4 6 8-2

-1.5

-1

-0.5

0

0.5

1

1.5

2


Motor Current [A]

Fo

rce

[N

]

V

= <5 m/s

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

-2 0 2 4 6 8-0.5

0

0.5

1

1.5

2

Sub-Vehicle Empirical Thrust (VB = 7.4)

Motor Current [A]

T [N

]

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

133


SV Subsystem Analysis


PowerFLOWComputational Fluid

Dynamics (CFD) simulation INPUTS:

Item Symbol ValueSV Wing Area S 0.447 m2

SV Chord Length c 0.271 mSV Wingspan b 0.6858 m

SV Mass m 0.380 kgFree-stream Airspeed V 15 m/s

Free-stream Air Density 0.98 kg/m3

Free-stream Air Viscosity µ 18.27e-6 Pa*sAngle of Attack -8 to +22 (2 discrete)

Reynolds Number Rec 218046

Item Symbol ValueZero-lift Angle of Attack L0 -1.351

Zero-Lift Drag Coefficient D0 0.01781Lift Slope CL 0.0278 per

Stall Angle of Attack stall 18-20Unity L/W Angle of Attack ULW 1.6565

Maximum L/D (L/D)max 5.5934Zero-Lift Pitch Moment CM0 -7.85e-3

Pitch Moment Slope ∂CM/∂ -0.0048 per

PowerFLOWComputational Fluid

Dynamics (CFD) simulation OUTPUTS:

134


Wind Tunnel Testing


0 0.5 1 1.5 2 2.5 3 3.5 40

1

2

3

4

5

6

7

8

9

10

11Voltage-Current Calibration Curve

Voltage [V]

Cu

rre

nt [A

]

Data

y = 3.3517V+0.10383

ESC Max, 3-cell

ESC Max, 2-cell

1 1.5 2 2.5 3 3.52000

3000

4000

5000

6000

7000

8000

9000

10000Propeller Speed as a function of Propeller Pulse Width

Pulse Width [ms]

Pro

pe

lle

r S

pe

ed

[rp

m]

Trial 1, 15 m/s

Trial 2, <5m/s

Exponential Fit

135


SV Level-2 EBD



0.3 0.35 0.4 0.45 0.5 0.55 0.60

2

4

6

8

Required Motor Current and Maximum Motor CurrentC

L = 0.19075 at = 6

o, t = 25.2min

mSV

[kg]

I M [A

mp

]

Ireq

CB = 2.1 A*hr

20 25 30 35

0.4

0.5

0.6

0.7

Allowable SV Mass for Varying EnduranceC

L = 0.19075 at = 6

o, C

B = 2.1A*hr, V

B = 7.4V

Endurance [min]

mS

V [kg

]

SV Payload capacity


The battery is chosen by selecting two bounds on the motor current draw:1. High bound: Maximum allowable current from the power and energy balances2. Low bound: SV stall speed

Item Mass (kg)

Airframe 0.170

Servos (2) 0.032

Propeller 0.005

Electric Motor 0.048

CUPIC 0.030

TOTAL 0.296

Superfly Stock Mass Budget without Battery:

137


SV FBDs


tan2

1

0tan2

1

cos

0sin

0cos

max,max,

2

LDL

stall

L

x

x

CCS

WV

DWSCV

DT

TWLF

DTF

z

x

T

W

L

D

stall

stallD

L

stallC

C

WT

cossin,

max,

138


SV Power Analysis


tV

tPVCI

tPtVIVC

tPtPVCE

PPP

M

PICBBM

PICMMBB

PICMBBB

PICMB

3600max,

Power and Energy Balances

M

pmM

M

pmMM

IV

VIfT

VITVP

)(

Linear Thrust Model

0 0.5 1 1.5 2 2.5 30

1

2

3

4

5

6

Maximum Allowable Sub-Vehicle Motor Current Draw (treq

= 30min)

CB [Amp*hr]

I ma

x [A

]

Model

COTS Available

139


New SV Thrust Model


tV

tPVCI

tPtVIVC

tPtPVCE

PPP

M

PICBBM

PICMMBB

PICMBBB

PICMB

max,


• The thrust model is adjusted to the wind tunnel test data and is compared to the power and energy balances.

0 1 2 3 4 5 60

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

Sub-Vehicle Adjusted Linear Thrust Model (VB = 7.4)

IM

[Amp]

T [N

]

V

= 6 m/s

V

= 8 m/s

V

= 10 m/s

V

= 12 m/s

V

= 14 m/s

140


SV: Aerodynamics


-20 -10 0 10 20 30-1

-0.5

0

0.5

1

SV Aerodynamic Data at Re = 272819


Co

effic

ien

t

Lift

Drag

Moment

Stall

0 0.05 0.1 0.15 0.2 0.25-1

-0.5

0

0.5

1

SV Drag Polar at Re = 272819

CD

CL

• Key Aerodynamics•Stall AoA: 18°• L/Dmax: 5.59

141


SV: Aerodynamics


Conclusions:1. Superfly maximum L/D is 5.5934, typical of delta-wing aircraft2. L/W > 1 for all reasonable trim angles of attack3. Superfly can fly at 5 m/s (minimum required) for all angles of attack

-10 -5 0 5 10 15 20 25-6

-4

-2

0

2

4

6

8


L/D

, L

/W

Sub-Vehicle L/D and L/W

L/D

L/W

L/W = 1

-10 -5 0 5 10 15 20 250

2

4

6

8

10

12

[deg]

TR [

N]


V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

Superfly Max

142


SV CFD Inputs and Outputs


PowerFLOWComputational Fluid Dynamics (CFD) SV simulation INPUTS:

Item Symbol ValueSV Wing Area S 0.447 m2

SV Chord Length c 0.271 mSV Wingspan b 0.6858 m

SV Mass m 0.380 kgFree-stream Airspeed V 15 m/s

Free-stream Air Density 0.98 kg/m3

Free-stream Air Viscosity µ 18.27e-6 Pa*sAngle of Attack -8 to +22 (2 discrete)

Reynolds Number Rec 218046

Item Symbol ValueZero-lift Angle of Attack L0 -1.351

Zero-Lift Drag Coefficient D0 0.01781Lift Slope CL 0.0278 per

Stall Angle of Attack stall 18-20Unity L/W Angle of Attack ULW 1.6565

Maximum L/D (L/D)max 5.5934Zero-Lift Pitch Moment CM0 -7.85e-3

Pitch Moment Slope ∂CM/∂ -0.0048 per

PowerFLOWComputational Fluid Dynamics (CFD) SV

simulation OUTPUTS:

143


Sub-Vehicle CFD Data


-10 -5 0 5 10 15 20 25-6

-4

-2

0

2

4

6

8


L/D

, L

/W

Sub-Vehicle L/D and L/W

L/D

L/W

L/W = 1

-10 -5 0 5 10 15 20 250

2

4

6

8

10

12

[deg]

TR

[N

]


V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

Superfly Max

144


Wind Tunnel Testing


Item Testing RangeAngle of Attack 6 (+/-3)

Free-stream Airspeed 5-15 m/s (5 m/s discrete)Free-stream Air Density 0.98 kg/m3

Motor Voltages 7.4V; 11.1VMotor Current 0-6 Amps; 0-10 Amps

145


Wind Tunnel Testing


0 0.5 1 1.5 2 2.5 3 3.5 40

1

2

3

4

5

6

7

8

9

10

11

Voltage-Current Calibration Curve

Voltage [V]

Cu

rre

nt

[A]

Data

y = 3.3517V+0.10383

ESC Max, 3-cell

ESC Max, 2-cell

1 1.5 2 2.5 3 3.52000

3000

4000

5000

6000

7000

8000

9000

10000

Propeller Speed as a function of Propeller Pulse Width

Pulse Width [ms]

Pro

pe

lle

r S

pe

ed

[rp

m]

Trial 1, 15 m/s

Trial 2, <5m/s

Exponential Fit

146


Wind Tunnel Testing


-2 0 2 4 6 8 10-1.5

-1

-0.5

0

0.5

1

1.5

2

2.5

3


Motor Current [A]

Fo

rce

[N

]

V

= <5 m/s

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

-2 0 2 4 6 8 10-0.5

0

0.5

1

1.5

2

2.5

3

3.5

Sub-Vehicle Thrust (VB = 11.1)

Motor Current [A]

T [

N]

V

= 5 m/s

V

= 10 m/s

V

= 15 m/s

147




Voltage (V) Speed (m/s) Thrust (CFD)(N) Thrust (Data)(N) Percent Difference Empirical Current (Amp)

11.1 5 0.1945 0.1990 2.2713 0.4333

11.1 10 0.7779 0.7500 3.7227 1.7751

11.1 15 1.7503 1.6785 4.2789 4.3849

7.4 5 0.1945 0.1850 5.2144 0.7189

7.4 10 0.77790 0.7450 4.4189 2.3480

Wind tunnel testing results:

Conclusions:1. Wind tunnel test data shows a slightly non-linear trend for T = f(IM)2. T = f(IM) model may be adjusted to fit wind tunnel test data and include propulsion

system efficiencies3. Considerations:

1. CFD did not model the wind tunnel or modified SV airframe2. Uncertainty in wind tunnel force measurements: 0.2N

148


SV Flight Test


Frank Dilatush performed a flight test of the Superfly.• SV Mass:

• 425g• Power Supply:

• 11.1V, 1.32A*hr• Achieved Endurance:

• ~23 minutes• Approximate angle of attack:

• 2°-10°• Approximate Motor Current:

• 2A-5A

• The maximum current allowed from the power balance is lower than necessary!

• The estimated endurance is ~4 minutes less due to non-constant battery voltage.

0 0.1 0.2 0.3 0.4 0.5 0.6 0.71.5

2

2.5

3

3.5

4

Superfly Flight Test (VB = 11.1V)

CL = 0.51925 at = 6

o, t = 22.833min

mSV

[kg]

I M [

Am

p]

Ireq

CB = 1.32 A*hr

149


Sub-Vehicle Components


Component Mass (kg)CG

X-Location (cm)CG

Y-Location (cm)CG

Z-Location (cm)

Foam Body MOI known MOI known MOI known MOI known

Servo (x2) 15.6 1.3 16 0

Propeller 5.594 3.4 29.2 0

Servo Rod (x2) 6.5018 2.1 32 0

Motor Controller 19.2028 3.4 22.5 0

Motor 48.194 3.4 26 0

Battery 95 3.05 11.9 0

CUPIC 30 .5 25 0

DM Attachment Point 13 6.875 11 0

MOI of the Foam Body about SV CG • Ixx = 9.29186 × 10-5 kg-m2

• Iyy = .0030131826 kg-m2

• Izz = 0.003070775 kg-m2

MOI of the SV about SV CG • Ixx = .00231183 kg-m2

• Iyy = .00301318 kg-m2

• Izz = .00307078 kg-m2

150


SV Center of Gravity


+x

0 cm 55.88 cm

Foam Body29.831 cm108.516 g

Propeller29.2 cm5.594 g

Servo Rod (x2)32 cm

13.032 g

Speed Controller22.5 cm

19.2028 g

Servo (x2)16 cm31.2 g Motor

26.5 cm48.194 g

Battery11.9 cm

95 g

DM Attachment

11 cm??? g

CG of System22.3 cm

CU PIC??? cm

30 g

151


SV Center of Gravity


CG Location for DM of 8g:CGx = 22.3063cmCGz = 2.2658cm

Recommended CG:CGx = 24.13cm +/- 0.64cm

6 8 10 12 14 16 1824.5

25

25.5

26

26.5

27

27.5

28

28.5

Mass of DM Attachment (g)

Lo

ca

tio

n o

f C

U P

IC (

cm

fro

m th

e n

ose

)

Determination of CU PIC Location Based On DM Attachment Size

PIC Location

Beginning of Prop Cut-Out

Max Mass of DM Attachment = 13.64 g

152




tV

tPVCI

tPtVIVC

tPtPVCE

PPP

M

PICBBM

PICMMBB

PICMBBB

PICMB

3600max,


M

pmM

M

pmMM

IV

VIfT

VITVP

)(

Linear Thrust Model

0 1 2 3 4 5 6 7 8 9 100

0.5

1

1.5

2

2.5

3

3.5

4


total = 60%)

IM

[Amp]

T [

N]

V

= 6 m/s

V

= 8 m/s

V

= 10 m/s

V

= 12 m/s

V

= 14 m/s

0 0.5 1 1.5 2 2.5 30

1

2

3

4

5

6

Maximum Allowable Sub-Vehicle Motor Current Draw (treq

= 30min)

CB [Amp*hr]

I ma

x [

A]

Model

COTS Available

153


Wind Tunnel vs. CFD


• A dynamic thrust test was performed in the ITLL wind tunnel• Percent differences from 2 – 5% for thrust model based on CFD• Test validated CFD model

Voltage (V)

Speed (m/s)

Thrust (CFD)(N)

Thrust (Data)(N)

Percent Difference

Empirical Current (Amp)

11.1 5 0.1945 0.1990 2.2713 0.4333

11.1 10 0.7779 0.7500 3.7227 1.7751

11.1 15 1.7503 1.6785 4.2789 4.3849

7.4 5 0.1945 0.1850 5.2144 0.7189

7.4 10 0.77790 0.7450 4.4189 2.3480

Wind tunnel testing validated CFD model and improved thrust model.

154


SV Stall Speed


0.25 0.3 0.35 0.4 0.45 0.5 0.55 0.6 0.654.5

5

5.5

6

6.5

7

Stall Speed for varying SV MassC

L,max = 0.51925 at = 18

o

mSV

[kg]

Vs

tall [

m/s

]

155


Operating Temperatures


Item Operating Low (Celsius)

Operating High (Celsius)

Storage Low (Celsius)

Storage High (Celsius)

SV Battery 0 45 -6.7 65.5

SV Motor - 110 - -

CUPIC Rate Gyro -55 125 -65 150

CUPIC Microcontroller -40 125 -65 150

CUPIC Pressure Sensor -40 125 -40 125

CUPIC Xbee Module -40 85 - -

• Limiting temperatures are all for the SV Battery

156


SV Payload


Item ValueTransmission Frequency 2.4GHz

Mass 9g (camera + RX)Signal-to-Noise Ratio 48dB

Wireless Transmission Range 300m line-of-sightPixels 250,000

Voltage Supply 9VCurrent Draw 100mA

•Sub-Vehicle payload: Wireless RC MiniCam•Requires 9V power supply, m = 46g•Total payload mass is 54g (<87g)

•May be used for visual verification of deployment sequence•May be mounted in an arbitrary orientation

Helihobby.com

157



Primary VehicleElectrical Design

PV Level-2 EBD


EDEPV Electrical

Design


System Architecture

Design Elements




Software Design Elements (SDE)

Software Drawing Tree


SDE

Detailed Communication Diagram


SDE


Sub-Vehicle Software Design

Stability & Controllability Analysis


• PowerFLOW was used to determine stability of the SV• Bill Pisano’s code includes a simulator of his autopilot in MatLab/Simulink

• Needs 3 stability derivatives• Cyβ = -0.0922• Cmβ = 0.0091• Cnr = -0.0018

SDE

-150 -100 -50 0 50 100 150 200 250-150

-100

-50

0

50

100

150

X Position [m]

Y P

ositio

n [m

]

Simulated Sub-Vehicle Loiter Path (rLoiter

= 100m, t = 150s)

Flight Path

Initial Position

End Position

-120 -100 -80 -60 -40 -20 0-100

-50

0

50

100


Real Axis

Ima

gin

ary

Axis

Short-Period

Phugoid

-1.5 -1 -0.5 0 0.5 1 1.5 2 2.5 3 3.5-20

-10

0

10

20


Real Axis

Ima

gin

ary

Axis

Spiral Divergence

Roll Subsidence

Dutch Roll

Sub-VehicleCUPIC Software

Design

164


System Architecture

Design Elements





System Architecture

Design Elements




PV & SV Component Testing


Primary Vehicle Component Testing

Test # Description Req. # Verified

1.TST1 PV Endurance; static engine test: Verify that the PV can fly for 30 minutes with the chosen engine (15 minutes with a SF=2) in the worst case scenario (fully loaded with 4 SVs the entirety of the flight), and worst case drag. (A model of the system shall be given the worst case parameters to determine endurance) A static test of the engine and fuel system shall determine the longest the engine can be run while verifying fuel tank size need, type of engine chosen and fuel consumption.

0.SYS10.SYS41.subSYSa11.subSYSa21.subSYSa32.subSYSa2

1.TST2 PV Endurance; battery: Verify that the chosen batteries for the PV, under the worst case power draw scenario can last 30 minutes (15 minutes with a SF=2). A model of the system shall give the worst case parameters to use to test the endurance. (Batteries onboard the PV will control the CUPIC and actuators for the DM.)

0.SYS11.subSYSa111.subSYSb4

Sub-vehicle Component Testing


2.TST1 SV Mass: Verify the mass of the SV. This can be checked by massing every component added to the SV using a scale. Modeling can show how different SV total mass changes flight characteristics.

1.subSYSb61.subSYSc11.subSYSc2

2.TST2 SV Endurance / Power; battery test: Verify that the SV batteries can last 30 minutes (15 minutes with a SF=2). Test that the batteries can supply the CUPIC and other elements on board the SV with a constant voltage and amperage necessary for them to operate over a period of 30 minutes.

0.SYS21.subSYSc72.subSYSc2

2.TST3 SV motor test: Verify a current CFD model. A dynamic test of the SV motor and modified SV in the wind tunnel in order to get a thrust as a function of motor current in order to verify a CFD model.

0.SYS2


DM & C&DH Component Testing


Deployment Mechanism Component Testing


3.TST1 DM Release; bench test: Verify that when the PV receives a signal, it controls the servo that pulls the pin in the release mechanism. Bench testing the prototype by applying a drag force, a weight simulating the SV, and a moment. This test would measure the force required to manually pull the pin perpendicular to drag.

0.PRJ10.SYS6

3.TST2 DM Release On Demand; car test: Verify that the pin can be pulled while a mock-up SV is attached to the DM, traveling at approximately 15 m/s. Test the prototype by attaching the DS to a car and drive the car at 15 m/s (the drop velocity ~ 35 mph) and manually pull the pin.

0.PRJ11.subSYSb22.subSYSa3

Command and Data Handling Component Testing


4.TST1 GS receiving data; ground test: Verify that the GS can receive data from multiple CUPICs at different distances and that sending the "release" command would not interfere with data collection. Verify this with a ground test, with the CUPICs not integrated on any of the vehicles, at different locations from the GS.

0.SYS41.subSYSa4


Interface Testing


Interface TestingTest # Description Req. # Verified0.TST3 Interface, PV with DM; Material Modulus: Determine Young's modulus and Poisson's ratio of balsa and

plywood used in the PV to increase the accuracy of the ANSYS model using a 3 point bend test. Using force and displacement measurements the modulus and Poisson's ratio of the materials can be calculated.

0.PRJ22.subSYSb3

0.TST4 Interface, PV with DS; flight characteristics: Verify with that the PowerFlow / MATLAB stability models are accurate; Using a full-systems flight test, measure takeoff distances and climb rates.

0.SYS5

0.TST5 Interface, PV with DS; stability: Verify with that the PowerFlow / MATLAB stability models are accurate; The stability derivatives from PowerFlow (the linearized longitudinal and lateral models) will be verified through flight testing where the elevator/rudder/aileron control deflections are recorded from the RC controller (pulse rates will map what deflection the servo give) and a 6 degree IMU is used to measure roll rates, etc; all of this information will then be plugged back into MATLAB to verify that how the full system flies, and what the model outputs are the same.

0.SYS5

0.TST6 Interface, PV with DM; Structural Integrity; Unmodified: Verify that the DM will not decrease the structural integrity of the PV. A model created in ANSYS will be verified by implementing load tests on the strut attached to the PV without reinforcement to the original fuselage.

1.subSYSb5

0.TST7 Interface, PV with DM; Structural Integrity; Modified: Verify that the DM will not decrease the structural integrity of the PV. A model created in ANSYS will be analyzed and then verified by load tests on the strut attached to the PV with reinforcement to the fuselage.

1.subSYSb52.subSYSb22.subSYSb3

3.TST3 DM Release On Demand; Vibe test: Verify that when the PV receives a signal, it controls the actuator that pulls the pin in the release mechanism at different frequencies. Test the prototype by attaching the DM, strut and SV to a vibration table and shaking the system at different frequencies that were determined by the ANSYS and PowerFlow models.

0.PRJ12.subSYSa3

3.TST4 DM Release On Demand; car test; verify drop model: Verify that when the PV receives a signal, the actuator pulls the pin and the SV is released. Test the prototype by attaching the DM, strut and SV to a car; drive the car at a velocity of approximately 15 m/s (the drop velocity ~ 35 mph) and send the command to the actuator to release.

0.PRJ12.subSYSa3


Fully-Assembled Static Testing


Fully-Assembled System Static Testing


0.TST1 Verify functionality of the CUPIC: Ground Test: Test the CUPIC to confirm that it can collect GPS data and stream that data to the GS at different ranges, up to a 2 mile radius. This will be ground tested by walking the CUPIC (not integrated with the system) in different directions, up to 2 miles from the GS in an open area (away from buildings) for a minimum of 30 minutes (15 minutes with a SF of 2).

0.SYS41.SYS21.SYS31.subSYSa121.subSYSd81.subSYSd91.subSYSd112.subSYSb1

0.TST8 Full-system Mass: Verify the mass of the entire system. This can be checked by massing every component added to the PV, and the PV itself.

1.subSYSa51.subSYSb12.subSYSa1

3.TST5 DM Release On Demand; Static full-assembly test: Verify that when the PV receives a signal using the CUPIC, it controls the actuator to pull the pin and release the SV. Static test with the DM integrated to the PV and SV in order to verify functionality and tolerance of mounting and functionality of the CUPIC and actuator interface.

0.PRJ11.subSYSc81.subSYSc91.subSYSc102.subSYSa3


Fully-Assembled Flight Testing


Fully-Assembled System Flight Testing

Test # Description Req. # Verified0.TST2 Verify functionality of the CUPIC: Flight Test: Test the CUPIC to confirm that it can collect GPS data and stream that data to the GS at different

ranges up to a 2 mile radius. The CUPIC functionality test will have a final verification during a full system flight test.0.SYS41.SYS21.SYS31.subSYSa121.subSYSc61.subSYSc111.subSYSd11.subSYSd2

1.subSYSd41.subSYSd51.subSYSd61.subSYSd71.subSYSd101.subSYSd121.subSYSd132.subSYSb1

0.TST9 Operating Flight Environment: Verify that the current environment is allowable for flight. This means verifying that temperature is within the range of 0˚C to 40˚C, no moisture in the air, and wind speeds less than 7 m/s.

0.SYS81.subSYSa101.subSYSc52.subSYSa42.subSYSa52.subSYSc1

1.TST3 PV Endurance; flight test: Verify that the PV can fly for 30 minutes (15 minutes with a SF=2) in the worst case scenario (fully loaded with 4 SVs the entirety of the flight). A fully-assembled system flight test, all 4 SVs attached for 30 minutes without deployment, shall be flown to test endurance.

0.SYS11.subSYSa71.subSYSa8

1.TST4 PV Stable Flight: Verify that the PV is in stable flight before deploying SVs. This will be determined by measuring the change in any direction (X,Y,Z) over time. Stable flight is defined as a maximum allowed change of TBD in any direction over TBD time. It is also verified by the Pilot rating scale. This experiment will be used to verify the PowerFlow modeling.

1.subSYSa61.subSYSa8

1.TST5 PV Velocity during Deployment of SV: Verify that the velocity of the PV when deploying the SV is at 15 m/s, +/- 0.01 m/s. (The velocity of the PV was determined using a "Drop Model" in MATLAB.) The velocity of the PV during deployment of the SV will be measured during full system flight tests.

2.subSYSa3

2.TST4 SV Stability: Determine the aerodynamic characteristics and stability of the SV. Then verify AVL and PowerFlow models with this data. This will be verified by remote controlling the aircraft and doing a similar test to used to verify system stability models: The stability derivatives from PowerFlow (the linearized longitudinal and lateral models) will be verified through flight testing where the elevon control deflections are recorded from the RC controller (pulse rates will map what deflection the servo give) and a 6 degree IMU is used to measure roll rates, etc; all of this information will then be plugged back into MATLAB to verify that how the SV flies, and what the model outputs are the same.

1.SYS1

2.TST5 SV Waypoint and Loiter: Verify that the SV is capable of flying to the waypoint and loitering for the remainder of the 15 minimum total flight times by the control of the CUPIC. This will be tested during a flight test to ensure that the CUPIC and clock on board are capable of meeting this requirement.

0.SYS31.subSYSc31.subSYSc41.subSYSd3

2.TST6 SV Endurance; flight test: Verify that the SV can fly for 30 minutes (15 minutes with a SF=2) using a flight test. During this test, the functional SV will be dropped from the PV and expected to fly for a total of 30 minutes.

0.SYS2

3.TST6 DM Release On Demand; flight test: Verify that when the PV receives a signal, the CUPIC controls the actuator to pull the pin and release the SV. The functionality of the DM release on demand will have final verification during full system flight testing.

0.PRJ11.subSYSb32.subSYSa3

4.TST2 GS receiving data; flight test: Verify that the GS can receive data from multiple CUPICs at different distances using a flight test. Also verify hat sending the "release" command would not interfere with data collection.

0.SYS41.subSYSa41.subSYSd6


System Tests


Ref MeasurementsRequired measurement fidelity and

durationRequired Equipment Notes/Comments

0.TST1 Transmission rateReceiving rateCollect GPS data and compute velocity, acceleration and distance data.

Duration: 30 minutesUp to a 2 mile radius from GSMeasurement Fidelity: 0.1 m resolution

CUPIC,Xbee GS,At least 2 team members (1 to walk around with the CUPIC, and 1 to monitor the GS)A program that calculates the velocity and acceleration

* MUST SUCESSFULLY TEST BEFORE INTEGRATION TO PV OR SV *(Verify the transmission rate)

0.TST2 Transmission rateReceiving rateCollect GPS data and compute velocity, acceleration and distance data.


Full system: PV, SV, GS * FULL SYSTEM FLIGHT TEST *This test will verify the ability to use the CUPIC on the PV and SVs as well as the interface with the GS.

0.TST3 Stress (Force applied over the cross-sectional area)Strain (displacement over the original length of the material)

Resolution of the load cell: +/-2NDuration: 4 times per trial, 16 total tests

(#) Balsa wood pieces dimensions: (#) Plywood pieces dimensions: Instron machine with a 50kN load cell3 point break test set-up

This test will determine the modulus of the two materials that comprise the PV structure. This information will be used in the ANSYS model for further analysis.

0.TST4 Takeoff distanceClimb rate

Velocity resolution: +/- 0.1 m/s Distance resolution: +/- 0.1 m

GPS data from CUPICMeter tape

* FULL SYSTEM FLIGHT TEST *

Compare takeoff distance and climb rates recorded to the predicted behavior in Travis' endurance code


System Tests




0.TST5 Pulse RatesRollPitchYaw

Accuracy requirements determined from model by TBD.Talk to Dr. G about stability to find the X acceleration resolution that may be needed then check the accuracy of the IMU to determine the feasibility of this measurement

6 Degree IMUFully-assembled system: PV, SV, GSOscilloscope

Plot the models vs. information collected to see how they line up and to verify the models

0.TST6 Forces appliedLocation of forces applied with respect to the fuselageMoment appliedStrain & Stress

Duration: TBD minutesFidelity: +/- TBD N

Loads will be worse-case scenarioPV and DM strut integrated w/o reinforcementStrain gauges

* MUST TEST BEFORE MOVING ON TO SIMILAR TEST WITH REINFORCEMENT *Worst case loading scenario needs to be used along with strain gauges to verify the model

0.TST7 Forces appliedLocation of forces applied with respect to the fuselageMoment appliedStrain & Stress

Duration: TBD minutesFidelity: +/- TBD N

Loads will be worse-case scenarioPV and DM strut integrated w/o reinforcementStrain gauges

* MUST SUCESSFULLY TEST BEFORE INTEGRATION OF DM TO PV *Worst case loading scenario needs to be used along with strain gauges to verify the model

0.TST8 Mass of each component added to the PV Mass resolution: 0.01 g Scale and components Modeling can show how different system total masses changes flight characteristics.

0.TST9 TemperatureHumidityWind speed

Duration: 30 minutes (flight time)Temperature Res.: +/- 2 degrees CHumidity Res.: +/- 5 %Wind speed Res: +/- 0.1 m/s

TBDWeather station that is used for the rocket lab could be a possible resource depending on the need for ASEN 2004

* FULL SYSTEM FLIGHT TEST *Environmental measurements can be used to determine how much of an effect the environment has on the full-systems' performance.


PV Sub-System Tests




1.TST1 Specific fuel consumption (g/kN*s)ThrustTimeTorqueRPM

Duration: 30 minutes (full flight time)Load cell resolution: +/- 2 NFuel consumption res.: 0.01gTime resolution: 0.01 sDynamometer res: TBD

Load cellDynamometerScaleTimer

Used to verify Travis' endurance code. Used to verify specifications of the motor and make endurance models more accurate if necessary.

1.TST2 Battery capacityTime it would take to drain the battery with a constant worst case power draw

Duration: 30 minutes Time res: +/- 1 secondCapacity res: TBD

Batteryresistors (to draw power instead of using the components, otherwise use CUPIC, servos, and high-speed cameras)power regulator

* MUST SUCCESSFULLY TEST BEFORE INTEGRATION TO THE PV *Battery capacity over time; compare to models

1.TST3 Measure remaining fuel (mass of full tank- mass of tank after flight)Total Flight Time

Duration: 30 minutes Time res: +/- 1 secondFuel remainder res: +/- 0.01 kg

Fully-assembled systemScaleTimer / Clock

* FULL SYSTEM FLIGHT TEST w/o SV DEPLOYMENT *Compare results with Endurance code

1.TST4 Change in XChange in YChange in ZTimeVelocity

Duration: 30 minutes (full flight time)Change in direction res.: +/- 0.1 m resolutionTime res: +/- 1 secondVelocity res: +/- 0.1 m/s

Fully-assembled SystemGS


Compare with stability models done in PowerFlow. Plot change in directions over time.Plot velocity with respect to time.

1.TST5 Velocity of PVTime SV takes to drop away from the PV

Time res: +/- 1 secondVelocity res: +/- 0.1 m/s

Fully-assembled System * FULL SYSTEM FLIGHT TEST *

The velocity of the PV at SV release can be put into the Drop model and then used to analyze the drop of the SV.


SV Sub-System Tests




2.TST1 Mass of each component Mass resolution: 0.01 g Scalecomponents

In order to make this a performance test, take the data and show how different SV total mass changes flight characteristics and verify through modeling.

2.TST2 Battery capacityTime it would take to drain the battery with a constant worst case power draw

Duration: 30 minutes Time resolution: +/- 1 secondCapacity resolution: TBD

CUPICSensorsBatteryConnections

* MUST SUCCESSFULLY TEST BEFORE INTEGRATION TO THE SV *

2.TST3 Varied windspeedVaried current going into the batteryNet force in the velocity direction

*** ASK JEFF WindtunnelModified SV assembled with motor and battery

This test was completed with the help from Professor Lawrence

2.TST4 Pulse RatesRollPitchYaw

Accuracy reqs determined from model by TBD.

6 Degree IMUFully-assembled system: PV, SV, GSOscilloscope

* SV FLIGHT TEST *Plot the models vs. information collected to see how they line up and to verify the models

2.TST5 Collect GPS data and measure how far off from the given point the SV is.

Duration: 30 minutes GPS resolution: +/- 0.1 m resolution / second

Fully integrated SVGS

* SV FLIGHT TEST **Dependent on CUPIC functionality tests under the Systems tab*

2.TST6 Flight TimeVelocityRemaining battery capacityTemperature

Duration: 30 minutesTime resolution: +/- 1 secondVelocity resolution: +/- 0.1 m/sBatery capacity resolution: TBDTemperature resolution: +/- 2 degrees C

Fully-assembled SystemGSTemperature sensor

* FULL SYSTEM FLIGHT TEST *This test will not be necessarily conducted by the autopilot. The SV may be controlled by an RC pilot for a portion of the flight time.


DM Sub-System Tests




3.TST1 Forces appliedMoments appliedForce to pull pin

Placement res: +/- 0.01 mForce res: Accuracy of sping scale? **ASK JASON

Spring scaleDM prototypeC-clamps

Simulated "worst case" aerodynamic loadsSimulated "worst case" friction environmentLift applied = 12NDrag applied = 12NCalculated torques

3.TST2 AccelerationsDisplacementsForce to pull pinTorsionStressStrain

Duration: TBDDisplacement res: TBDForce res: TBDAcceleration res: TBD

AccelerometersTesting beamSVDMPickup Truck

Trying to simulate flight environment.

3.TST3 AccelerationsDisplacementsForce to pull pinTorsionStressStrainHigh-speed VideoFrequency

Duration: TBDDisplacement res: TBDForce res: TBDAcceleration res.: TBDFrequency res.: TBDHight-speed video res.:TBD

AccelerometersTesting beamSVDMVibration tableStrobe LightLampsHigh-speed cameraMemory card and computer

Vibration test requested by Koster in an RFA from PDR.

3.TST4 AccelerationsDisplacementsForce to pull pinTorsionStressStrain

Duration: TBDDisplacement res: TBDForce res: TBDAcceleration: TBD

AccelerometersTesting beamSVDMPickup TruckLaptop

Verify and compare with drop model.

3.TST5 Measure placement Placement res: +/- 0.01 mAngle: TBD

Verifies C&DH on the ground.

3.TST6 Time to drop out of the way of the PVPath taken when released

Duration: TBD Time: +/- 0.1 secondResolution of Camera: TBD

Fully-assembled system(Including high speed cameras)


Compare with drop model.


C&DH Sub-System Tests




4.TST1 Measure transmission rateMeasure distance from GS


CUPIC,Xbee GS,At least 2 team members (1 to walk around with the CUPIC, and 1 to monitor the GS)A program that calculates the velocity and acceleration

* TESTED ALSO UNDER SYSTEMS FOR CUPIC FUNCTIONALITY -- TESTS GO HAND IN HAND *

* MUST SUCESSFULLY TEST BEFORE INTEGRATION TO PV OR SV *

(Verify the transmission rate)

4.TST2 Measure transmission rateMeasure distance from GS


Full system: PV, SV, GS * FULL SYSTEM FLIGHT TEST *

* TESTED ALSO UNDER SYSTEMS FOR CUPIC FUNCTIONALITY -- TESTS GO HAND IN HAND *



System Architecture

Design Elements




Detailed Mass Budgets


Mass Budget

Product Mass (kg) # Total (kg)Notes

Primary Vehicle

Airframe 5.500 1 5.500

Engine 1.520 1 1.520

Ignition Battery 0.120 1 0.120

Propeller 0.000

Fuel Tank 0.000

Fuel 0.320 1 0.320

Control Surface Servo 0.047 4 0.186

Throttle Servo 0.037 1 0.037

Receiver 0.010 1 0.010

Control System Battery 0.120 1 0.120

CUPIC 0.030 1 0.030

Wiring/Cables/Mountings/Etc 0.100 1 0.100Margin

Subtotal 7.943

Sub Vehicle

Airframe 0.110 4 0.440

Motor 0.048 4 0.193

Motor Controller 0.019 4 0.077

Propeller 0.006 4 0.022

Servo 0.016 8 0.125

Servo Connection Rod 0.006 8 0.048

Battery 0.095 4 0.380

CUPIC 0.030 4 0.120

DM Attachment 0.016 4 0.064

Payload 0.100 4 0.400

Subtotal 0.446 1.869


Upper Mounting Rod 0.195 1 0.195

Upper Mounting Reinforcement 0.026 2 0.052

Lower Mounting Rod 0.200 1 0.200

Lower Mounting Reinforcement 0.046 2 0.092

Lower Mounting Connector 0.020 1 0.020

Mechanism 0.040 4 0.160


Subtotal 0.795

Total 10.607

Available 12.670 30 min Endurance

Balance/Additional Margin 2.063 16.3%

179

PMP

Detailed Fiscal Budget


Cost Estimates

Product Cost ($) # Total ($)Notes

Primary Vehicle

Aircraft

Provided by RECUV

SIG Rascal 110 Blue/White ARF 1.20-1.50, 110" 319.99 2 639.98

APC 16x8 Pattern Propeller 6.47 4 25.88

Servos/Electronics

Futaba NR4F 4.8V 1500mAh Receiver NiCd Flat 27.43 2 54.86

Futaba S3305 Servo High-Torque Standard w/Metal Gears 27.43 8 219.44

Futaba S3004 Standard Ball Bearing Servo 10.57 2 21.14

Hitec Aileron Extension 24" 4.51 4 18.04

Hitec Aileron Extension 36" 5.47 2 10.94

Hitec Y-Harness 6" 4.49 2 8.98

EMS Heavy Duty ESC Switch J 11.99 4 47.96

Transmitter/Receiver

Futaba 6EX 6-Channel 2.4GHz Transmitter/Receiver 156.79 1 156.79

Futaba R617FS 7-Channel 2.4GHz FASST Receiver 73.49 1 73.49

Engines

G26 Electronic Ignition 299.99 2 599.98

Futaba NR4F 4.8V 1500mAh Receiver NiCd Flat 27.43 2 54.86

Fuel System

Dubro Gas Conversion Stopper 1.45 2 2.9

Dubro In-Line Fuel Filter Large Scale 3.29 2 6.58

Dubro ID Tygon Tubing 1/8" 3' 4.49 4 17.96

Dubro Fuel Line Clips Large (4) 1.99 3 5.97

MPI Aluminum Fuel Dot 3.99 2 7.98

Slimline M4 Gas System w/Container 38.49 1 38.49

Miscellaneous

Hobbico Latex Foam Rubber 1/4" 3.39 3 10.17

Tower Hobbies Digital LCD Mini-Tachometer 15.67 1 15.67

CUPIC

CUPIC 400.00 1 400.00

Subtotal 2438.06

RECUV Expense 2038.06

MADS Expense $ 400.00

Sub Vehicle

Airframe 50.00 4 200.00

Motor 15.00 1 15.00

Propeller 1.50 1 1.50

Servo 20.00 2 40.00

Battery 35.00 1 35.00

CUPIC 400.00 1 400.00

Subtotal $ 691.50


Mounting Bracket 22.50 1 22.50

Mini Nylon Steel Pushrod 2.50 4 10.00

Aluminum T6061 0.375"x6' Rod 11.00 2.0 22.00


Servo Extension 7.00 4 28.00

Subtotal $ 334.50

Ground Station

RC Transmitter 0.00 1 0.00Provided by RECUV (See PV)

XBee Pro Ground Station 140.00 1 140.00

ICD2 Programmer 0.00 1 0.00Available from ITLL

Laptop 0.00 1 0.00Team Supplied

Subtotal $ 140.00

Margin (30%) $ 1,200.00

Total $ 2,766.00

Available $ 4,000.00

Balance/Additional Margin $ 1,234.00 30.9%

180

PMP

Spring Schedule 1


PMP

Spring Schedule 2


PMP

Documents

Critical Design Review - colorado.edu Pisano’s UPI autopilot simulator •Stability derivatives calculated from PowerFLOW System Architecture Deployment Mechanics ... Critical Design