sae 2015 final report

SAE AERO DESIGN WEST – 2015

Van Nuys, California

Design Report By

TEAM - 310

MICRO CLASS

B.M.S COLLEGE OF ENGINEERING

BANGALORE , india

B.M.S College of Engineering, Team-310

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Table of contents

List of figures and tables………………………………………………………………………….4

1.0 Executive summary

1.1 System overview……………………………………………………………………………..5

1.2 Competition projections……………………………………………………………………..5

1.3 Innovations…………………………………………………………………………………....5

2.0 Schedule summary………………………………………………………………………....6

3.0 References……………………………………………………………………………………6

4.0 Loads and Assumptions

4.1 Standard Environmental assumptions…………………………………………………….7

4.2 Landing Shock……………………………………………………………………………….7

5.0 Design Progress

5.1 Conceptual Design and Initial Weight Estimate …………………………………………8

5.2 Preliminary Design………………………………………………………………………….10

5.2.1 Configuration selection…………………………………………………………………..10

5.2.2 Airfoil Selection……………………………………………………………………………10

5.2.3 Planform Selection………………………………………………………………………..11

5.2.4 Detailed Design……………………………………………………………………………12

5.3 Geometry Selection And Optimization………………………………………………….....12

5.4.1 Servo Sizing………………………………………………………………………………..13

5.4.2 Propulsion…………………………………………………………………………………..13

5.4.3 Competitive Scoring and Strategy Analysis……………………………………….……14

6.0 Analysis

6.1 FEM Analysis of Wing Spar……………………………………………………..…………..14

6.2 CFD Analysis of 2D Airfoil………………………………………………………………..….16

6.2 Performance Analysis

6.2.1 Drag Analysis………………………………………………………………………………..17

6.2.2 Propulsion…………………………………………………………………………………....18

6.2.3 Wind Tunnel Test…………………………………………………………………………....18

6.3 Stability………………………………………………………………………………………….20

6.4 Thrust Analysis…………………………………………………………………………………21

7.0 Assembly And Sub-Assembly

7.1 Testing…………………………………………………………………….……………………...23

7.1.1 Aerodynamics………………………………………………………………………………....23

7.1.2 Dynamics and Controls………………….........................................................................23

7.1.3 Structures………………………………………………………………………………………23

7.1.4 Propulsion………………………………………………………………………………………23


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7.1.5 Flight Testing………………………………………………………………………………….23

8.0 Manufacturing

8.1 Overview………………………………………………………………………………………....24

8.1.1 Laser Cutting…………………………………………………………………………………..24

8.1.2 Fuselage………………………………………………………………………………………..25

8.1.3 Wing…………………………………………………...........................................................25

8.1.4 Tail………………………………………………………………………………………………26

9.0 Conclusion

10.0 Appendix

List of figures and tables

FIGURE NO. FIGURE NAME PAGE NO.

1 Milestone Chart 6

2 Average Temperature in 2014 7

3 Weather Forecast 7

4 Landing Shock 7

5 Basic Design Progress 8

6 Optimization Flowchart 9

7 Mass Indexing of Fuselage 9

8 2D Analysis of High Lift air foils 11

9 S1223 Air foil 11

10 Airfoil Analysis 11

11 Geometry Selection And Optimization Flowchart 12

12 Matlab Code for Competitive Scoring 14

13 Von Mises Stress on deformed wing with respect to unloaded condition 15

14 Von Mises stress on Deformed wing 15

15 Shear Stress induced on Wing 16

16 Deformation Induced on Wing 16

17 CFD Analysis showcasing Surface Velocity 16

18 CFD Analysis showcasing Pressure Distribution 16

19 Virtual Wind Tunnel Test 19

20 Laser Cutting of Airfoil 25

21 Airfoil Section 25

22 Prototype Plane 26

TABLE NO. TITLE PG. NO.

1 Figure of Merit 11

2 Dimensions Table 13

3 Drag Analysis Table 18

4 Thrust Analysis Table 22

5 Master Test Schedule 25


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1.1 System Overview This report presents the design, manufacturing, and testing processes conducted by Team

Yaksha of B.M.S College of Engineering in preparation for the SAE AERO WEST 2015 at Van

Nuys, California, USA. The winner of this competition is the team that achieves the highest

overall score, which consists of three parts, the average of successful flight round scores,

operational availability points depending on payload fraction of the aircraft and the assembly

demonstration bonus. To achieve this task in the micro-class competition, the aircraft designed

by this team utilized a conventional aircraft configuration, which was modified to excel at

providing heavy lift capabilities. The conventional aircraft configuration was chosen over more

advanced designs because of its performance predictability and ease of construction. In the

following sections, we outline the design process used to create our micro aircraft.

Calculations used to justify the aerodynamic, propulsion, and stability performance are

detailed and a final weight build-up is given. Finally we present the model diagram of the

aircraft based on the above considerations.

1.2 Competition Projections

Our team performed an analysis to identify how to maximize score based on the key

requirements. From this analysis, the team determined that minimizing the empty weight and

maximising the payload fraction are the main factors for maximizing the total flight score. This

was accomplished by maximizing the aerodynamic efficiency of the aircraft. The team after

several brainstorming sessions decided to choose a conventional wing design because of its

proven reliability, high stability and ease of construction. A high lift to drag ratio would help

increase the system efficiency. A rectangular carbon fibre fuselage was chosen consisting of

4 rods for its high strength to weight ratio. Truss structures were cut to minimize the net weight

of the aircraft. The aircraft uses a tractor propulsion system consisting of an AXI brushless

out-runner motor, 1000 mAh Li-Po battery pack and a 9x4.5 inch propeller.

1.3 Innovation The team tried innovative methods in fabricating the aircraft. Since the micro class has belly

landing instead of a landing gear mechanism most of the fuselage body was made of carbon

fibre. The team also used laser cutting techniques to fabricate the wing. Rapid prototyping

machine Protocentre 999 was used to make the model and test it on the wind tunnel. Hinges

were added at the tail to accommodate the aircraft inside the prescribed container comfortably.


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2.0 SCHEDULE SUMMARY

The team constructed a rigid timeline with effective time distribution and was able to execute

them in regular intervals. Our team was divided into several section like design team ;

Fabrication Team ; Electronics ; Report and Red Team. The outline chart of our work processis

shown below in the following Gantt Chart.

3.0 REFERENCES

1. UIUC Applied Aerodynamics Group, UIUC Airfoil Coordinates Database.

www.ae.illinois.edu/m-selig/ads/coord_database.html

2. Carpenter and Houghton, “Aerodynamics for Engineering students”, Butterworth

andHeinemann, 17March 2003.

3. Raymer, Daniel P. 1999 “Aircraft design, A conceptual approach”, AIAA Alexander Bell

Drive and Reston. October 15, 2009.

4. WB Garner, Model Airplane Propellers. March 2009

http://www.dc- rc.org/pdf/Model%20Propellers%20Article.pdf

5. <http://www.csd.net/~cgadd/eflight/calcs_servo.htm

6. Selig, Michael S. “High-Lift Low Reynolds Number Airfoil Design”. Published in Journal of

Aircraft, Vol 34, No. 1, January-February 1997

7. Wilcox, David C (1998). "Turbulence Modeling for CFD". Second edition. Anaheim: DCW Industries, 1998. pp. 174.

Figure 1. Milestone Chart

http://www.ae.illinois.edu/m-selig/ads/coord_database.html


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4.0 Loads and Assumptions

4.1 Standard Environmental Assumptions:

The standard sea level conditions are shown above and these are taken into consideration

while designing the preliminary design. The team conducted a literature survey on the weather

conditions at Van Nuys; California. The MET department reports were studied and it was found

that the temperatures stay between 25 degrees Celsius (max) and 10 degrees (min). The

winds may range from 9 to 11 km/hr which comes under the category of light breeze under

Beaufort Scale. Though these are not alarming weather conditions; the team left no stone

unturned in prepare for the worst. The predicted weather during the period of the competition

are shown below collected from the survey done.

Figure 3. Weather forecast

4.2 Landing Shock

Calculating the landing shock of

the aircraft is an important factor

of analysis when there is no

landing gear present in the micro

class division. Thus a MATLAB

program was created to analyse

the force experienced by the belly

of the fuselage as it has to land

on the fuselage in the flight

rounds. A plot of Force vs Initial impact time is shown to understand the amount of force

that’s going to be acting at the impact time.

Figure 2. Average temperature in 2014

Figure 4: Landing Shock


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5.0 Design Process

The final flight score in this year’s competition is a cumulative of the average of successful

flight round scores, operational availability points depending on payload fraction of the aircraft

and the assembly demonstration bonus. So to attain a good flight score, the plane should be

capable of carrying a very high payload having the least empty weight possible. To keep the

requirements of the competition and constraints in view an initial weight estimate was done.

The empty weight without electronics was estimated to be 300g and the total aircraft weight

estimated to be around 500g. Over the years it has been seen that the maximum payload

fraction achieved by top performing teams in micro class division has been around 0.7. Thus

a payload fraction of 0.7 was set up as target. Our conceptual design steps are outlined in the

following flow diagram.

5.1. Conceptual Design and Initial Weight Estimate

After a few brainstorming sessions the team decided to go with a conventional rectangular wing design. The factors instrumental behind our conceptual design are size, lifting capacity, wing loading, ease-of-design and construction, and drag effects and also the new introduction of the shape of container from this year in micro class division. Based on the size of plane, the speeds at which it would fly, a high lift, low Reynolds number airfoil (S1223) was selected by the team would be best suited for the competition. According to the literature survey performed, it was found out that an elliptical wing is highly efficient. But due its complicated construction a straight rectangular planform with no

Figure 5: Basic Design Process


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taper ratio was selected for the simplicity and ease of fabrication. A conventional tail was chosen for adding stability to the design. The main idea was to make the dead weight of the aircraft as minimum as possible. Thus a single carbon fibre rod is used to connect the main part of the fuselage and the tail. Moreover truss structures are used to optimise the weight to strength ratio as possible. The following flowchart describes the thought process implemented while doing the initial weight estimate.

Figure 6: Optimisation Flowchart

Figure 7: Mass Indexing of Fuselage


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Next, we performed a mass indexing program to have an estimate whether actually the

proposed aircraft weight matches the calculated weight. The above diagram shows the mass

indexing of the entire fuselage part. It shows that with the available electronics it stands at

around 240g.

5.2 Preliminary Design:

5.2.1 Configuration Selection:

The configuration selection is the most important consideration of the entire project. The team

studied various configurations to determine which would suit the problem statement and uplift

the fight score and came up with 3 configurations namely, Monoplane (conventional), Biplane,

Flying wing. The various configurations were ranked points based on the literature survey

made by the team. Figure of merit (FOM) was used to decide, with the maximum of 5 points

allotted to each parameter.

Parameters Weightage Conventional Biplane Flying wing

Ease of Construction

0.6 4 3 3.5

Cost 0.3 4 3 3.5

Empty Weight 0.7 3 3 4

Stability and control

0.8 4.5 4 3.5

Historical data and literature

survey

0.6 4 3 2

TOTAL 11.7 9.8 9.95

Table No.1: Figure of Merit

Thus it can be found out from the above table that the conventional plane configuration is

giving us a better trade-off than the other designs thought of. The Biplane and the Flying Wing

configuration were eliminated because of its complexity in construction. Hence we concluded

that the conventional configuration was best suited for the competition requirements because

of high ‘Cl’ value at given Reynolds number, the ease of construction, historical data available,

structural simplicity and its cost involved.

5.2.2 Airfoil Selection:

The team conducted study on a wide range of airfoils. Among many airfoils suitable for high

lift application. E423, S1223, S1210 airfoils were analysed using XFLR5 software. It was

concluded that S1223 airfoil is best suited for payload lifting applications at a lower Reynolds

number. It has highest CL value at zero angle of attack which is required for high payloads and

also it has maximum lift coefficient. This airfoil has also proven to be effective for these

applications in the past. The primary reference for our final selection was Selig’s paper on the


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S1223, which showed its superiority in comparison with other heavy lift airfoils. We also

referenced World of Krauss for low Reynolds number data.

Figure 8. 2D Analysis of high lift airfoils

Figure 9.S1223 airfoil

A batch analysis of the 2-D wing was setup on XFLR5 analysing the lift coefficient, drag

coefficient, moment coefficient for varying angle of attack and Reynold’s number for laminar

flow. The yielding data was then validated through Wind tunnel testing of a 3:8 scaled model.

Figure 10.S1223 Airfoil Analysis


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5.2.3 Planform Selection

Though a rectangular planform is more efficient, a rectangular planform was selected due to

the lack of team’s past fabrication experience. A taper ratio of 1 was used to further simplify

the design calculations and construction. For an empty weight of 1.1 lb and a payload fraction

of 0.67, the required lift produced by the wing should be 3.51lb. Flight velocity of 36 ft/s was

assumed based on historical data. Using XFLR-5 software initial 2D-lift coefficient was found

to be 1.5. Using the formula for the lift, the wing area required for the corresponding lift was

found to be 1.59 ft2(0.148 m2).

5.2.4 DETAILED DESIGN

3-D Lift Coefficient (From finite wing theory):

The 2-D coefficients are for wings with an infinite aspect ratio. Wings of finite aspect ratio i.e.,

3 -D wings need a correction to the lift coefficient. The lift coefficient of finite wing is lower than

the infinite aspect ratio wing, due to lift induced drag which manifests as wing-tip vortices

reducing the overall lift coefficient. The finite wing lift curve has the same zero lift angle of

attack, and pivots about this point, clockwise. The reduced slope 𝑑𝐶𝑙𝑑∝ derived from theory

is given by,

From the above equations and, 𝑑𝐶𝑙𝑑∝ equals 0.08 Coefficient of lift for finite aspect ratio is given by the equation 𝐶𝐿= 𝑑𝐶l𝑑 𝛼∝+𝐶𝐿∝=0

And 𝐶𝐿∝=0 = 1.34. This gives a corrected 𝐶𝐿 value of 1.385 at ∝=5° A 3-D wing analysis using Vortex Lattice Method in XFLR-5 software resulted in a CL value

of 1.33 which closely matches with the theoretical value calculated above. A 3-D wing

analysis using Vortex Lattice Method in XFLR-5 software resulted in a CL value of 1.33

which closely matches with the theoretical value calculated above.

5.3 Geometry Selection and Optimization

The geometry of the aircraft was selected and optimized using Excel spreadsheet and

MATLAB program which incorporates take off analysis, stability analysis and also the

constraints put forth by the competition. The program was constructed as per the flowchart

shown in the Figure.


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Figure 11. Geometry Selection and Optimisation Flowchart

Structural Component Dimensions

Total Length 34.929 inches

Fuselage Length 24.101 inches

HT Chord 4.72 inches

HT Span 11.22 inches

VT tip Chord 4.560 inches

VT Span 5.118 inches

Wing Span 55.110 inches

Wing Chord 5.5 inches

Wing Area 2.1035 ft2

HT Area 0.644 ft2

VT Area 0.1292 ft2

Aspect Ratio 10.22

Table No. 2: Dimensions Table

5.4.1 SERVO SIZING

Maximum Torque required for our constraints is given by the following equation:-

Torque required = 8.5 E-6 [C2 V2 L sin (S1) tan (S1) / tan (S2)]

C = Control surface chord in cm, L = Control surface length in cm, V = Speed in MPH,

S1 = Max control surface deflection in degrees, S2 = Max servo deflection in degrees,

For ideal performance and efficient manoeuvring, the chords of the control surfaces are taken to be one third of the main stabilizers chords. From this, the dimensions of the control surfaces are obtained as: Elevator =3.5 cm ×28.5cm and Rudder = 3.5cm x 12cm.The performance characteristics suggest a minimum take-off speed of 11 m/sec. The maximum torque required to operate the elevator and rudder at the above speed are 0.67 kg-cm which is well below the value of the maximum torque of the selected servos which is 2 kg-cm


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5.4.2 Propulsion

The propulsion system was designed to produce the required thrust to overcome drag at cruise

while weighing as little as possible. It was designed to operate for a minimum of two minutes

to accommodate for multiple circuits of the course at maximum speed plus a safety factor. The

most important of these priorities was to produce enough power and, in turn, thrust to achieve

flight.

To get an initial estimate of the power output needed, we used the rule of thumb of

approximately 30 W per one lbf of weight lifted. At the beginning of the design process, our

initial weight was approximately 1.8kg and thus our required power was approximately 120 W.

Given the power estimate we determined a range of battery specifications that had the

capability of producing a minimum of 120 W. In order to achieve this power output, batteries.

As such, a three cell battery was chosen. Given the flight duration of two minutes, a maximum

amp-hour value for the battery was calculated to be 850 mAh at an operational current

between 20A and 25A and battery specification voltage of 11.1V.

5.4.3 Competitive Scoring and Strategy Analysis

Figure 12: MATLAB Code for competitive scoring

For analysing the flight score and to maximise the points based on the competition scenario

at the airfield; a simple program was constructed on MATLAB which is shown above. This

program is constructed on assumptions that the length of the container must be 56 inches

and a bonus of 1.7. Moreover all the flight rounds must have successful flight round scores.

On giving input of around 6 payload fractions we determine how the score is varying or

depending on the payload fractions given. So on input [0.3 0.42 0.45 0.5 0.52 0.55] we get a

total cumulative score of 120.2907 points. This is figure which we want to improve upon and

trials are going on to determine the maximum scoring possible.


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6.0 ANALYSIS

6.1 FEM ANALYSIS OF WING SPAR

CFD results are directly analogous to wind tunnels results got in a research facility they both

speaks to sets of information for given stream arrangements at distinctive Mach numbers,

Reynolds number, and so forth. What's more mulling over is that CFD permits perception of

stream properties without exasperating the stream itself, which is not generally conceivable

with ordinary measuring instruments.

In real word scenario, the impinging air parameters are difficult to record, thus, hybrid

initialization is used. For zero angle of attack the speed forms will be symmetrical and in

accordance to rule of Bernoulli’s. In a symmetrical aerofoil at no occurrence, the circulation of

speed and in this way the pressure along both surfaces would have been precisely the same,

cancelling off one another to a subsequent aggregate lift power of zero.

The wing structure experience various types of loads amid every period of the flight which

incorporates take-off, ascension, voyage, arriving, and landing. We are interested to find out

to stresses acting on the wings. The wing geometries created were then meshed. The sizing

option used was proximity and curvature. In additional to body sizing, edge sizing and face

mapping was carried out. The geometry was then subjected to stacking state of 40 N of

dynamic load, the load was connected via containing the central ribs that will accompany

contact with the airframe. This was the supposition taken, i.e. an example of a cantilever bar.

Figure 13. Von Mises Stress on the deformed wing with respect to unloaded condition

The stress analysis of the wind indicates an anhedral of 10 for a wing loading of 40N. This

effect is counteracted by manufacturing the wing rib with carbon fibre reinforcements. The

anhedral is also corrected by the Lift generated during flight which in turn gives the wing a

dihedral of 20.


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Figure 14: Von Mises Stress on the deformed wing

Figure 15: Shear Stress induced on wing

Figure 16: Deformation induced in the wing


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6.2 CFD Analysis of 2D Airfoil

The above figures show the CFD analysis is done at a free stream velocity of 10 m/s and a Reynolds number of 100,000 for Sellig-1223 2-D airfoil.

6.2 PERFORMANCE ANALYSIS

6.2.1 Drag Analysis The drag analysis of the plane is carried out in three parts-wing drag, fuselage drag and the tail drag. The team estimated the total drag co efficient of the wing using component build up method. The total drag coefficient of the airplane is the sum of these three drag components. CD=CDO+CDI where CDi= CL

2 / πeAR CDO=Cdo + Cd misc Cdo =1Σn (FF. Cf .Swet/Sref) Evaluation of Cd0: Swet=Sexp [1.977+0.52 (t/c)] FF of wing and tail = [1+0.6/( 𝑥/𝑐)2∗(𝑡/𝑐)+100 (𝑡/𝑐)4 ]* [1.34 M0.18 cos(𝛬𝑚)0.28] FF of fuselage= 1+𝑓0/𝑓3+𝑓/400 where f= 𝑙/𝑑… ..eqn 6.17 Cf = 1.328/ (Rcut-off)0.5 Rcut-off = 38.21(𝑙/𝑘)1.053

WING FUSELAGE VERTICAL TAIL

HORIZONTAL TAIL

Sref (mt2) 0.1918 0.066642 0.0115 0.03135

Swet (mt2) 0.379882 0.0497923 0.232584 0.060846

Swet /Sref 1.9806 0.74 2.022 1.94

Using the above equation, CDO is found for each component in the Table as shown below.

Figure 17: CFD Analysis showing surface velocity Figure 18: CFD analysis showing pressure

distribution


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WING FUSELAGE VERTICAL TAIL HORIZONTAL TAIL

FF 1.327 1.15 1.1123 1.1097

l/k 35897.43 16117.948 2617.4 2724.87

Rcutoff 206896.3189

1029139.447 151772.9065 154628.81

Cf*10-3 2.919 1.31 3.4 3.37

Table No.3 Drag Analysis Table

Using the formula for the total drag, we have 𝐶𝐷=𝐶𝐷𝑂+𝑘𝐶𝐿2+𝑘′(𝐶𝐿−𝐶𝐿𝑚𝑖𝑛𝐷)

CDo = (𝐹𝐹)( 𝐶f)(𝑆wet/𝑆ref) k =1/(𝐴𝑅)𝑒 =0.039

k’ (for blunt rectangular plan form wings)=0.02

𝑘 CL 2+𝑘′(𝐶𝐿−𝐶𝐿𝑚𝑖𝑛𝐷) = 0.0728

For fuselage; FF= 1 + 60/𝑓3 + f/400 = 1.15

Therefore, CDo = n=1Σ4 [(𝐹𝐹)( 𝐶f)(𝑆wet/𝑆ref)]

CDo = 0.0236 ,Cmisc. = 0.0728 , CDi = 0.039 CL 2

Therefore, CD = (0.0964) + 0.039CL 2 Thus the drag polar curve is given below.

Drag Polar curve

6.2.2 Propulsion

The propulsion system was so designed to provide the necessary thrust to overcome the drag

experienced during flight. The initially designed system had a high power motor producing a

thrust of 3.96 lbs with a 11x5.5 propeller and required a higher capacity (mAh) battery which

invariably increased the weight of the system above the optimum level. Hence the team

optimized the plane design further in order to reduce the drag so that a lower thrust rated

motor can be used to provide the required outcome. Finally the team selected a motor that

provided 3.3 lbs thrust and having 1380 Rpm/volt. This motor required a 40A ESC and a lower

capacity battery to provide the same flight time necessary for a single course at max speed in

the competition still weighing less than the previously designed system.


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In order to meet the required power output, batteries of different cell counts were considered.

For a given flight duration of two minutes, a 3 cell battery providing 11.1v, having a capacity

of 1000 mAh rated to have a constant 45c discharge rate and a burst of 90c was chosen. The

motor selected required a max current of 35A/60s from a 3 cell Li-Po along with a

9”x4.5” propeller for the designed system specifications to meet the necessary static thrust

for our application.

6.2.3 Wind tunnel Test

Data Validation was done through testing a scaled model in the wind tunnel facility at the

university. A 3:8 scaled model of the designed aircraft was fabricated using Rapid prototyping

technique to fit the dimensions of the test section of the wind tunnel (0.5mX0.5mX1m). A

blockage ratio of 60% was assumed due to high aspect ratio. Experimental data was gathered

using a 6 component strain gauge and the Coefficient of lift and drag were calculated for

various angle of attack.

Graph of lift coefficient V/S angle of attack

-10

0

10

20

30

40

50

60

-10 -5 0 5 10 15

CL/CD

-0.5

0

0.5

1

1.5

2

2.5

-10 -5 0 5 10 15

Lift Coefficient (CL)


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Graph of CL/Cd v/S angle of attack

It was found that the values

were obtained to an

accuracy of approximately

80% of theoretical results.

A virtual wind tunnel was

conducted to find out the

drag coefficients.

6.3 Stability

Horizontal tail:

The horizontal tail volume co-efficient is given by

Vh was found to be 0.45, and the wing area, chord being – 308inch2 and 5.5inch

respectively, we find the optimum tail area to be 34.54inch2. Taking tip width to root width

ratio to be 0.45, we find the dimensions of the horizontal tail to be:

Tip width: 4’’; Root width: 2.4’’; Span: 11.2’’;

Vertical Tail:

The vertical tail volume co-efficient is given by

Vv was found to be 0.03, and the wing area, wing span being – 308inch2 and

56inch respectively, we find the optimum tail area to be 23.44inch2.

The vertical tail area is thus taken as 24inch2.

35% of the area of the tail area was considered for the elevator and rudder. With regard to

the above calculations an inverted flat plate was selected since it could satisfy all the

moment balancing requirements arising from the pitch moment generated by the main wing

and thus optimum stability was obtained.

Vh * Sw * c = St * lt

Vh=Tail Co-efficient

Sw=Wing area

St=Tail area

C=Mean average chord

Vv * Sw * b = St * lt Vh=Tail Co-efficient

Sw=Wing area

St=Tail area

b=Wing Span

Figure 19. Virtual Wind tunnel test


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6.4 Thrust analysis

With respect to the nature of power output of propeller motors and the inherent relationship

between a motor and a propeller, it was necessary to consider motor choice and propeller

sizing in tandem. As such, the first general relation to consider was thrust versus RPM. Next

to be considered was propeller pitch versus thrust. Consequently, lower pitches meant higher

thrust values, but also lower RPM. Therefore, a higher power motor was required to spin a

propeller with less pitch at the same RPM as one with more pitch.

Finally, the relationship that brought the voltage of the battery and RPM of the motor together

was the KV parameter of the motor. KV is a measure of the RPM per Volt a motor can output,

and in doing so, how much torque a motor outputs. Lower K meant lower RPM with more

thrust which was typical of slow flying planes, much like the very aircraft our team was

designing [10]. Using these general guidelines as well as the parameters of the battery, the

motor Kv value was selected to be 1390 Kv. To determine prop size, it was necessary to relate

thrust and power to pitch speed, Vpitch. Pitch speed can be calculated using equations :

𝑇 = 𝑃/𝑉𝑝𝑖𝑡𝑐ℎ

Aprop = T/(0.5*p*Vc 2)

Dprop = 2*(Aprop/π)0.5

Dynamic thrust had to be greater than static thrust. It is given by

T = TC ρ V2D2

Since static thrust is dependent on the diameter of the propeller disk and the outgoing

airspeed, we determined that a nine-inch diameter, 4.5 inch pitch propeller was needed for

our application.

RPM/V 1380

Diameter x Pitch 9 x 4.5

Thrust 1500g

Table No.4 Thrust Analysis Table


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7.0 ASSEMBLY AND SUBASSEMBLY

The aircraft was designed to be a four fixture plane consisting of two parts of the wing and two

parts fuselage. The larger structural parts of the aircraft had to be assembled in a certain order.

The fuselage was first assembled by bolting the carbon fibre rods to the mounting plates in a

rectangular fashion. Another carbon fibre boom of diameter 12 mm was fixed to the middle

plate which stretches to the tail. Then the horizontal stabilizer is fixed to the boom followed

by the vertical stabilizer.

The team fixed the control surfaces to the appropriate positions using hinges, then in the next

step electronics were integrated into the body. Then the assembly of the electronic

components starts with the motor which is fixed to the front mounting plate and secured

using screws. To ensure correct orientation and fit of the components special care was taken.

Connecting wires were cut into required length to optimize the space required to fit in all the

electronic components. The Electronic speed controller was then connected to the motor and

secured by applying glue. The battery is then placed in the appropriate position and secured

using zip ties as needed to be removable. The receiver is then placed suitable and the

connections between the ESC and the receiver are made.

For the wing assembly first the laser cut ribs of the airfoil shape were placed at the required

location on a wide sheet of balsa. Then a 5mm carbon fibre rod was made to pass through

the ribs and glued at the appropriate locations. After the wing is fabricated, the structure of the

wing is covered using monocot. Ailerons are the attached to the wing structure and secured

using hinges.

Wing rib Manufacturing

Next the two servos for the ailerons are fixed and secured in place, the servos are then

connected to the radio receiver in order to facilitate it’s working. Simultaneously the servos

assigned for the rudder and the elevator are also placed and connected to the receiver.

Propeller is then fixed to the motor using the mounting unit. In the final step of assembly the

pre prepared cargo bay unit is fixed and secured to the fuselage in order to provide the

required space to carry the designated payload. Next the fuselage is covered with monocot.

The wing is then placed on the fuselage and firmly attached using nylon hinges and fibre tape.

Finally the plane was completely sealed and all the electronic components were checked for

flight readiness.


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7.1 TESTING

Objectives

After the prototype was complete, a series of tests were performed to evaluate the actual

performance of the aircraft. The prototype was tested in five categories: aerodynamics,

dynamics and controls, propulsions and structures. The goal of these tests was to verify

theoretical calculations and validate them with experimental data. Flight tests were also

performed to evaluate the overall aircraft performance.

7.1.1 Aerodynamics

Tests were run using the Wind Tunnel at BMS College of Engineering, Bangalore of

dimensions 0.5 m by 0.5m by 1 m. The main objective was to measure aerodynamic

performance by obtaining overall lift and drag coefficients at varying angles of attack and

free stream velocities from the measured aerodynamic forces. Lift and drag magnitudes were

then used for the propulsion analysis. All experiments were performed using the wind tunnel.

The aerodynamic coefficients tested in the tunnel were within 80% of the theoretical results.

7.1.2 Dynamics and Controls

The objectives of the stability tests were to ensure the aircraft had sufficient yaw, pitch, and

roll stability. Two tests were performed to measure the stability of the aircraft. In the first test,

wing size was adjusted to minimize drag while maintaining acceptable yaw stability. In the

second test, the elevators were tested to ensure adequate pitch and roll authority were

achieved while keeping a minimal overall weight. All tests utilized the wind tunnel apparatus.

7.1.3 Structures

Loading tests were performed on key aircraft structures to verify that failure loads exceeded

design requirements. The wing structure was tested by fixing the carbon fibre wing spar in the

Universal Testing Machine .The loads were applied were corresponding to a factor of safety

of 2 .The spar was able to withstand the applied loads with minimal deflection. These tests

identified potential failure points and design revisions were made accordingly.

7.1.4 Propulsion

A series of propulsion tests were executed to evaluate the performance of various motor,

battery, and propeller combinations. A wind tunnel and force balance were used to measure

the thrust produced at various flight speeds. A simple setup was fabricated on which the

motor was mount, and thrust was measured at different RPMs of the selected motors.

7.1.5 Flight Testing

Initially a full scale glider was constructed and glide tested to check the air worthiness of the

design. Later a full powered prototype aircraft was constructed and flight tested with different

payloads in order to test the stability and maximum payload capability of the aircraft.

The first few test flights resulted in frequent crashes due to stability and control inaccuracies.

These errors were a result of inaccurate calculation of CG. During landing of the next fight,

we witnessed another failure in the rib at the wing tip when the plane landed on its side. During

the next set of flights the ESC failed due to excess draw of current from the motor. Once these

issues were resolved by altering the position of the center of gravity and replacing the ESC

of higher current rating, a successful test flight was achieved.

A Master Test Schedule was constructed to analyse the different test results below.


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Test area Test Objective Team members delegated

Duration

Aerodynamics Wind tunnel Lift, drag, pitching

Abhiram, Sachin

15/11/14 – 22/11/14

Dynamics and control

Elevator, Ailerons, Rudder sizing

Minimum sizing for maximum control

Indrajeet, Sachin, Achyut

29/11/14 – 15/12/14

Structures Spar testing Spar dimensions for 20N load

Abhiram, Srinath, Abijith

29/11/14 - 2/12/14

Wing testing Wing loading and torsional strength

Indrajeet, Blesson

15/12/14 – 31/12/14

Propulsion Propeller Correct sizing Apratim, Siddesh

1/1/15 – 5/1/15

Motor Necessary thrust

Siddesh, Apratim

5/1/15 – 15/1/15

Battery Various sizes for thrust alterations

Adarsh, Vishwas

8/1/15 – 28/1/15

Flight test Flight data Voltage, amperage, wattage

Akshay, Adarsh, Indrajeet

30/1/15 – 15/2/15

Flight characteristics

Necessary design changes and optimization

Abhishek, Ayush, Kushal, Deevyendu

15/2/15 – 28/2/15

Table No. 5- Master Test Schedule

8.0 Manufacturing

8.1 Overview

The team after conducted a literature survey on the material and found that the carbon fibre

rods have higher strength to weight ratio than other conventional materials. The team

decided to use birch wood for rib sections. The material used for the spars in the wing is

reinforced carbon fibre rods. Balsa sheet covering and cap strips are used to maintain the

airfoil shape.

Since it is a two piece wing, each segment was connected using reinforced carbon fiber rods

running through the ribs having holes of 5mm diameter. Monokote is used for covering the

wing. Hinges were used to fix the ailerons. The parts were bonded together using

Cyanoacrylate adhesive (743 Adhesive).

The fuselage is a rectangular structure, built using 4 carbon fibre rods and three mounting

plates. Then a carbon fibre rod boom is connected from the second plate till the end where

the tail is mounted.


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8.1.1 Laser Cutting

Standard process for the laser printer was to model all parts in

software, produce a drawing with only the outlines of the parts.

The parts were printed from CATIA modelling software. For the

final model of the aircraft, the laser drafting cutter was

also used for the precise cutting and measurement of the

thick Monocot . The main challenges to using the laser cutter

were determining proper settings for each thickness of wood

and providing tolerance to structurally significant cut outs.

Mock printing sessions were conducted by the team prior to

cutting plane parts in order to discover the best printing settings

for different parts. Balsa wood and Monocot are soft materials

so they cut at a faster speed then the harder, more common

acrylic materials that the laser cutter is usually used for. Thus careful preventive measures

were undertaken. The ability to laser cut every part of the aircraft, whether it was Monocot,

balsa, or acrylic.

The main construction of the three parts of the aircraft is given below:

8.1.2 Fuselage

The main rectangular structure was constructed using four reinforced carbon fibre rods.

Two mounting plates were chosen and held together by screws. The rods were passed

through the third mounting plate in the middle. A 11mm carbon fibre rod was then mounted

from the middle plate. These rods provide the main strength and the rigid form of the entire

structure. The motor was mounted on the front plate of the fuselage

8.1.3 Wing

Balsa wood was used for the construction of the wing. It has a span of 1.4 metres. Birch wood

was used for rib sections at the ends of the wing to provide strength to the wing. Reinforced

Carbon Rods were passed through the 5mm holes to provide rigid form to the wing. Spars are

passed through 2mm holes which will take bending load stress. Each rib is placed after 2

inches. A prototype wing was built at first and tested. It was found to be heavy and stood

about 150g. Subsequent changes were made in lightening the overall weight. So the model

was optimized to the extent possible.

Figure 21: Airfoil Section

Figure 20. Laser Cutting Of Airfoils


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8.1.4 Tail

The conventional inverted-T tail design was chosen and the balsa wood was laser cut

according to design and truss structures were introduced similar to the wing. Control surfaces

were laser cut too and attached to the tail and the wings using a hinge mechanism. Next a

hinge was introduced at the horizontal surface to fit the tail within the prescribed diameter.

This has proved to be a major challenge and the team is currently working on 2 to 3 methods

and choose the best of them.

The first prototype of the plane is shown below:

Figure 22: Prototype Plane

Conclusion

SAE Aero design competition is an excellent platform for the upcoming engineers to get a

deep insight into the design and fabrication aspects of aero-modelling since it requires

extensive multidisciplinary work which in turn helps the students to develop team building

skills. This competition provides a practical hands-on experience to the team members in both

building and flying of the aircraft.

TEAM YAKSHA of BMS College Of Engineering, India has done a complete study and made

slight modifications on the conventional design, performed through engineering analysis and

has constructed a final model that meets the requirements laid out by the Society of

Automotive Engineers for the Aero Design West competition 2015.


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The test flights served as the basis for most of our modifications such as adjusting the centre

of gravity, increasing the stability. Working on the S1223 airfoil was a very challenging task as

the trailing edge of this particular airfoil is very narrow and hence too delicate, giving raise to

many complexities. Securing the electronic components in the hollow fuselage was an uphill

task.

It was a great learning curve of the young undergraduate team worked with passion and

dedication towards achieving the goal. Overall the experience gained while working towards

this competition proved to be very helpful and constructive and gave us an opportunity to have

a deeper look into the field of aero design and avionics.

Appendix A: Weight and balance data

Parts list with Weight data


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Appendix B: PAYLOAD PREDICTION GRAPH

Payload Prediction Graph

Appendix C: Model of the aircraft

Isometric View of the model

Payload = -0.0001(density altitude) + 2.9066

0

0.5

1

1.5

2

2.5

3

0 2000 4000 6000 8000 10000 12000

Pay

load

s (l

bs)

Density Altitude(ft)

Payload prediction


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sae 2015 final report

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