fall final design presentation

7/30/2019 Fall Final Design Presentation

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Dimitrios ArnaoutisAlessandro Cuomo

Gustavo Krupa Jordan Taligoski

David Williams

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Project Specifications

Airplane guidelines Flight guidelines ScoringProject Design Wing Tail BoomCalculations Performance Weight Fuselage sizing

Length, Width, Height Drag

Material Selection Aircraft PayloadCost AnalysisSummary

SummaryProduct SpecificationsDesign

Wing Tail BoomConcepts & SelectionMaterialsCalculations? Need or replacewith Concept analysis?Concept Analysis

Aerodynamics Stability & Control

Performance Structural Analysis Estimated Weight CostsManufacturing ProcedureEnvironment, Health &Safety


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Maximum combinedlength, width, andheight of

No more than fifty fivepounds ( ) with

payload and fuel. Single unmodified O.S

61FX

3

if flowninto No Fly zones x2

Multiple passes of field isallowed No touch and go

landings


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5


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Standard

Flying Wing

Minimalist

Canard Bi-Plane


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SelectionCriteria

Weight Rating WeighedScoreRating Weighed

ScoreRating Weighed

ScoreRating Weighed

ScoreRating Weighed

Score

PotentialLift 20% 7 1.4 9 1.8 8 1.6 8 1.6 7 1.4

PotentialDrag 10% 4 0.4 8 0.8 9 0.9 2 0.2 3 0.3

Durability 15% 9 1.35 5 0.75 3 0.45 7 1.05 7 1.05

Cost 10% 5 0.5 5 0.5 8 0.8 3 0.3 4 0.4

Ease of Manufactu

re5% 5 0.25 6 0.3 8 0.4 4 0.2 4 0.2

PotentialFlightScore

40% 8 3.2 6 2.4 7 2.8 7 2.8 7 2.8

100% 6.55 6.95 6.15 6.15


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Commonly used incommercial passenger aircraftas cargo area

Design Flush with fuselageStrength:

Good torsion resistanceWeight:

Heavier weight in

comparison to otheroptions of tail booms. 8

http://www.me.mtu.edu/saeaero/images/IMG_1215.JPG

Used in model aircraftand small helicoptersDesign:

Best done with carbonfiber (not permitted)

Strength:Low torsion resistance

Weight: Lightest weight design Design:

Greatly affects fuselagedesign

Strength: Great torsion resistance High stabilityWeight:

Highest weight comparedto other booms


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Drag 0.20 3 2 1Ease of Build 0.10 5 3 2

Maneuverability 0.15 3 4 5

Stability 0.35 4 4 5

Weight 0.20 4 4 3

Total 1.00 3.5 3.5

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Roots of both stabilizerattached to fuselageEffectiveness of verticaltail is large

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http://me-wserver.mecheng.strath.ac.uk/group2007/groupj/design/airframe/lower/image/conventionals.jpg


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FRP (fiber-reinforced plastics) not allowed

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http://cdn.dickblick.com/items/333/01/33301-8301-1-3ww-l.jpg

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With an approximated payload = 35 16 we can approximate the volume of thepayload based on densities of various commonmetals and their corresponding cost, and decide ona material for the payload.

From this analysis our payload will likely be Steel*Data selected from Callister 7 th edition

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Material Density (gm/cm 3) Cost (USD/kg) Volume (in 3) Cost (USD)Steel Alloy 7.85 0.5 123.414 7.94Stainless Alloy 8 2.15 121.1 34.13Gray Cast Iron 7.3 1.2 132.712 19.05

Copper Alloy 8.5 3.2 113.976 50.8


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CNC cutting for airfoil ribs, fuselage ribs, andstabilizersAs many as 3 prototypes in event of crashMost lightweight construction methodspossible


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14

Engine Magnum xls 61 1 $99

Balsa Wood Structure of aircraft, variouslengths and shapes ~50 ft. $100

Monokote Skin around structure ~50 sq. ft. $60

ServosControls flaps (elevator, aileron,

rudder, etc.) 5 $125Fuel Tank Holds fuel within fuselage 1 $5

Battery Powers servos and receiver 1 $15

Radio and receiverRadio controller for the plane

and the receiver to send control

functions to servos

1 $0

MiscellaneousItems

Wheels, pushrods, hardware,engine mounts, propeller TBD $75-$150

ShippingWill be Shipping supplies from

high fly hobbies located inDaytona Beach, FL

2-3 $14.95(per box)

Total *estimate *$509-$600


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-0.5

0

0.5

1

1.5

2

2.5

-20 -10 0 10 20 30

2D Lift Curve Re =3E+5 Stall Angle

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1.94 Design LC

Knowing the MTOW, we find

0

0.5

1

1.5

2

2.5

3

0 0.05 0.1 0.15 0.2

EPPLER 420

S1223RTL

UIR 1540

Airfoil data calculate for Cl_max

Lift Coefficient = 2.34Drag Coefficient = 0.048L/D = 48.8

Moment Coefficient = -0.202

According to the literature(Abbot), thevortex effects decrease 20% of the aircraft`slift coefficient.


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Wing span = 2.7 mRoot Chord = 0.32mTip Chord = 0.16 mM.A.C = 0.28 mTip Twist = - 2 degreesWing Area = 0.728 m^2Aspect Ratio = 10

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The software utilized was the Cea- VLM (vortex lattice method) Several iterations were made varying:

Wingspan Wing root and chord Taper ratio and its position

considering it s consequences to: Wing weight (estimated via the Cubic L aw) Wing lift and drag

this process was monitored by the: Oswald s factor


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-13

-11

-9

-7

-5

-3-1.4 -0.9 -0.4 0.1 0.6 1.1

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0

5

10

15

20

25

30

35

-1.5 -1 -0.5 0 0.5 1 1.5

The wing loads were estimatedutilizing the methodology proposed bySchrenk

In a later analysis this data will beused to size the wing spar by usingfinite element methods


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Initial elevator design Zero lift airfoil, 0 degree angle of attack Large pitching moment coefficient:-0.4296

Revised elevator design Zero lift airfoil, -9 degree angle of

attack Minimal pitching momentcoefficient: -0.0222

Also a negative lift airfoil canbe used


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OS 61 FX Suggested fuel tank cap: 350cc

12-13min flight Displacement: 9.95cc (0.607cu.in.) Bore : 24.0mm (0.945 in.) Stroke: 22.0mm (0.866 in.) Practical RPM: 2k~17k rpm Power output: 1.9 bhp @ 16k rpm Weight: 550g (19.42 oz.)

Deliver reliable and efficientpower to propel the aircraft. In the form of thrust with the help

of a propeller.


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Thrust is required topropel aircraft Requires energy (from

engine) to produce thrust

Force of thrust generatedby engine & propeller

Experimentally determine thrust:Thrust stand

Give accurate static thrust ratingsfor motor and propellercombinations


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The thrust-to-weight ratiois a fundamentalparameter for aircraftperformance Acceleration rates Climb rates

Max/min speeds Turn radiusHigher T/W will acceleratemore quickly, climb morerapidly and achieve highermax speed

Using a max take-off distance of 200 feet, areference T/W wascalculated,

=. .

. . . 0.204

The thrust required attake-off was calculatedusing Aximer TR = 5.83 lbf The thrust available attake-off is expressed by ,

TA =

TA = 15.19 lbf The aircraft will haveenough force to thrustthe 35 pound payloadinto flight.


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Assuming 85%efficiency of motorshaft power, the poweravailable is 1.615 hp.

The PR is importantwhen computing whatthe output needs to befor a given altitude andvelocity The motor performance is

fixed Other factors must be

adjusted to compensate

Converting tohorsepower yields avalue of 0.971hp The motor is sufficient

enough to create thrustfor the max payload of 35 pounds

=

= 0.204 47 55.71= 534.15


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Transfer mechanicalenergy from shaft intothrust.Propeller drag is a lossmechanism Robing engine of net

power outputthrust. Efficiency increases as

propeller size increasesRequires increased groundclearance and low tipspeeds.Optimize with diameter,pitch and blade count

Propellers can be sizedaccording to HP of theengine (2-blades eqn)

Results in 25 diameter Formula unsuitable forsmall scale RC

Propellersrecommended

Sport: 12x6-8, 13x6-7Aerobatic: 12x9-11

= 22 .


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Measuring various makesand models of propellerscould be useful. Build thrust stand

Recommended sportpropellers were analyzedwith ThrustHP Allows varying inputs of

propeller (diameter, pitch,blade count, make)

Approximate and recordthe RPM to reading close to1.9bhp *0.85=1.62bhp

Some useful outputs:Static thrust

=


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The thrust initially beginsat a large value butdecreases with increasingvelocity Weight and dynamic

pressure decrease

At cruise altitude thrustbecomes equal to weightthus, no additional thrustis needed to causemotionDrag tends to increasewith increasing velocitybecause the Reynoldsnumber is becomingmore turbulent yieldingmore drag effectively

0

1

2

3

4

5

6

7

0 50 100 150 200 250

T h r u s

t o r

D r a g - l

b s

Velocity - kts

Maximum & Cruise Speed

Total Thrust

Cruise Thrust

Drag


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The rate of climb (RC)is the rate at which anaircraft can safely andeffectively change

altitudesUsing Aximer thepredicted climb ratewith standard flightconditions at cruisevelocity was calculatedto be

RC = 12.543 ft/s

0

500

1000

0 50 100 150 200 250

C l i m

b -

f p m

Velocity - kts

Rate of Climb - Sea Level


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Performance ParametersClimb Angle 5.1670 degreesRate of Climb 0.1920 m/sVstall 10.6832 m/s

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0.00

10.00

20.00

30.00

40.00

50.00

0.00 5.00 10.00 15.00 20.00 25.00-2

-1

0

1

2

3

0 10 20 30

0

0.05

0.1

0.15

0.2

0.25

0.000 0.200 0.400 0.600 0.800 1.000 1.200 1.400

Components

C.G.

GeometricalConstraintAerodynamicalCenter

The components will be positionedaccording to the overall effect that they haveon C.G. The V-n Diagram gives an overview of theflight envelope by relating its velocities tothe load factor that the aircraft will undergounder that speed.


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= can assume a value of about 35.3 lbs which was themax payload of last year s 1 st place aircraft

can be determined using the following givens andrelations:

Given: fuel = 1.1371 g/cm 3 ; V tank 350 cm 3 ; g = 9.81 m/s 2

= fuel x V tank x g 3.904 N 0.8777 lbs can be estimated using a minimum ratio of 0.2 (W e/ Wo)

=

= .. .

45.22 lbs

= = 9.0423 lbs 55 lbs

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Utilizing a spreadsheet CG and Sizinganalyzer we were able to determine the sizingof the fuselage based on the wing dimensions

= 0.7*Wingspan = 6.20 ft Average diameter can be calculated using afineness ratio (FR) of 10 and the length of thefuselage

= =.

= 7.44 in (circular)If the cross section is noncircular, the heightand width can be attained using the relation,

=+ If we set H = 2W for clearance

purposesW = 4.96 in H = 9.92 in (rectangular)

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Wetted Area Estimation (blunt body)Circular Fuselage : = 2

12.682 ft 2Rectangular Fuselage : = 2

16.061 ft 2 Drag Estimation

= Assume : q = 1.0665 lb/ft 2 Re = 300,000 (laminar)

= 1.0665 12.682 .,

= 0.0328

= 1.0665 16.061.,

= 0.0415

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Magnum xl 61 engine uses 10% nitromethane ( 4CH3NO2 + 3O 2 4CO 2 + 6H 2O + 2N 2)Over the course of the semester it isestimated we will use a little over 4 gallons of nitro methaneThis translates to about 4 lbs of CO 2 greenhouse gas

The average passenger car produces thisamount in under 5 milesInsignificant amount of pollution


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Always keep fingers clear of a running engineWhen revving up, hold engine from verticalstabilizer, not behind engine or on wing leadingedge

Always refuel the aircraft in a well ventilated areaKeep fuel away from outside ignition sourcesAll members of team keep an eye on the flyingaircraft at all timesNever fly more than one plane at a timeWhen possible, wear hardhats when in the flyzone


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fall final design presentation

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