Nsit Sky Knights - Design Report(048) (1)

8/3/2019 Nsit Sky Knights - Design Report(048) (1)

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NSIT-SKYKNIGHTSSAE AERO-DESIGN FINAL REPORT

COLLEGE - NETAJI SUBHAS INSTITUTE OF TECHNOLOGY (INDIA)

TEAM NUMBER 048

TEAM MEMBERS:

1. ATHAK BHARADWAJ 424/IC/082. HIMANSHU SHARMA 443/IC/08

3. GAURAV KAUSHIK 436/IC/08

4. SUPRIYA TOMAR 510/IC/08

5. SWATI NEGI 511/IC/08

6. ISHA AGGARWAL 451/IC/09

7. CHITTESH SACHDEVA 619/MP/108. RAVI KAPOOR 652/MP/10

FACULTY ADVISOR : Prof. M.P.S. BHATIA


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ABSTRACT

The SAE Aero Design Competition is an international event which provides an exposure to

the undergraduate and graduate engineering students from many universities to the various

kinds of situations that engineers like us face in our real life work environment.

It challenges us to design, create, build and test a remote control airplane. Every feature

that comes into play when planning an aircraft right from the wing profile, dimensions,

center of gravity, materials etc. have to be designed from scratch.

During the design process we perform a series of studies and analysis in order to arrive to a

most favorable and desirable design solution.

The most favorable design solution will be one that will perform to the best of its abilities,

capable of lifting great loads, cost efficient while remaining as light as possible without

compromising the safety of the aircraft.

ABOUT REGULAR CLASS

Aero design features 3 classes of competition-Regular, Advance and Micro. Regular class

continues to be the class with the purpose to develop fundamental understanding of flight.

SAE also focuses on the importance of interpersonal communication skills of todays

engineer and therefore improve our written and oral communication skills SAE has devoted

a high percentage of a teams score solely to the design report and oral presentation in

front of the SAE judging panel.


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TABLE OF CONTENTS

1. SUMMARY 05

2. CONCEPT OF AERODYNAMICS 06

3. MATERIAL SELECTION 09

4. WING DESIGN AND CONFIGURATION 10

5. FUSELAGE DESIGN AND STRUCTURE 13

6. TAIL SELECTION 15

7. TAIL DESIGN 16

8. ENGINE 17

9. CONTROL MECHANISM 19

10. SERVO SELECTION 20

11. LANDING GEAR 21

12. DRAG ANALYSIS 25

13. PAYLOAD PREDICTED GRAPH 27

14. DIMENSIONS AND STRUCTURAL VIEW 28

15. TIMELINE 29

16. REFERENCES 30


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SUMMARY

Our report summarizes the team work of "NSIT SKYKNIGHTS", an aerial vehicle designed

by a student team of NETAJI SUBHAS INSTITUTE OF TECHNOLOGY for the SAE AERO

DESIGN EAST 2011.

This document portrays the design process, including some of the approaches that were

chosen by the team to achieve the objectives of the SAE, to create a radio controlled (RC)

aircraft that will lift the required weight, and not exceed the length, breadth and height

restrictions. Detailed analysis of the airplane will be presented, together with the technical

and experimental verification that justify the most relevant design considerations.

The main design prerequisites of the UAV were defined primitively in the design process i.e.

an aircraft which takes care of the vital prospects namely reduced drag, craft size

constraints, increased lift, structural integrity and stability.

Due to introduction of the new 1 minute payload loading and 1 minute payload unloading

demonstration the fact which was taken care of was the use of independent payload bay

assembly with easy weight loading access. Since use of some lightweight polyesters(fiber-

reinforced plastics) are restricted, the team will have to design a plane with using minimum

amount of material while supporting heavy loads in order to be at par with requirements.

Some of the fundamental decisions mentioned in this report are a repercussion of the

knowledge gathered from the participation of the SAE Aero Design team in previous

editions of the SAE.

Nonetheless, extensive studies were developed in order to eliminate some of the crucial

inefficiencies or to further improve some design offsets that proved to be unsuccessful.


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CONCEPT OF AERODYNAMICS

I.LIFT

Lift on a fixed wing aircraft is achieved by creating a difference between the pressures on

the upper surface of the wing and the lower surface of the wing. Notably by creating high

pressure on the lower surface and low pressure on the higher surface and by Bernoullis

principle. Bernoullis equation is the principle governing this pressure gradient.

The coefficient of lift, as a function of Aspect Ratio (AR) and angle of attack (), generally

follows the slope:

This relationship follows the linearity character for a limited range of angle of attacks.

Beyond that range, stall and inverted stall conditions create a non-linear relationship.

II.DRAG

Determination of drag characteristics was a more challenging task. Total aircraft drag

comprises of drag-due-to-lift, skin friction drag, pressure drag (aka form drag), and

interference drag from the combination of wing, fuselage, engine, tail, and landing gear.

The vortices created along the wing span traverse in the shape of spiral toward the

fuselage on the low pressure surface and towards the wing tip on the high pressure

surface. This interference drag is highest with a low-wing design and lowest with a high-

wing design.


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To estimate the total drag characteristics of the aircraft, all the aircraft components must be

considered. The basic expression for drag function can be obtained from Dr. Leland

Nicolais White Paper.

CD = CDmin + (CDduelift)*(CL - CLmin)2

CDmin is obtained from pressure and skin friction drag of all the aircraft components. Skin

friction drag is predominant for wing, fuselage and tail. The Reynolds numbers for these

components must be found to compute the coefficient of friction, which is computed as

outlined in Dr. NicolaisWhite Paper. The coefficient of friction can then be substituted into

the CDmin equation below. The Form Factor depends on the aircraft component being

analyzed.

The CD due lift factor replaces the K and K factors described in Dr. Nicolais method,

since 3-D airfoil data was generated from computer simulation.

AIRFOIL SELECTION

The selection of the airfoil of the estimated R/C aircraft (low speed high lift) depends upon

following factors; I. I. Airfoil drag, stall & pitching moment characteristics, thickness

available for the structure.

II. Substantial pressure differentials over a much greater percent of chord.

III. To maintain laminar flow over the greatest possible part of the airfoil.

To choose the airfoil for above characteristics, we came across following airfoils and

studied their properties in the domain of our requirements.


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From the above procured data of the airfoils, we zeroed upon to FX 63-137 airfoil as its

characteristics are best suited for the aircraft. As can be deduced from the above data, FX

63-137 has the best L/D (design lift coefficient) among the studied airfoils. Best L/D is the

point of the airfoil drag polar that is tangent to a line from the origin & closest to the vertical

axis. It has good lift properties with moderate drag parameters also does it have a high

negative pitching moment & convex pressure recovery.

FX 63-137


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MATERIAL SELECTION

we studied a wide range of materials to be used in construction of our aircraft which

includes mainly metals, plastics and wood. And thus categorized their properties

considering their density, weight and strength (stiffness, bending and compression).

Strength/weight(S/W) ratio required for wings and fuselage best overlapped on the range of

wooden materials. The strength properties of

woods under consideration are shown in the below

table. From the given data we concluded that balsa

wood is most appropriate for our use. As our stress

analysis of wings and fuselage had shown the

required strength and its weight constraints, so we

had to use materials with best available S/W ratio

as these components are required to be strong as well as lightest possible for better lift and

structural integrity in air. Balsa wood S/W ratio is much higher and favorable for the

fuselage and wings construction as it is the lightest wood with moderate strength (an

optimum combination of the two). Its stress analysis is shown below (solidworks 2009) to

study its strength under different types of applied forces. As the analysis portrayed, its

stiffness and compression stress lies with in a desirable range of strength. So we used

sheets, blocks and sticks made of balsawood. Sheets were used in fuselage and airfoil

construction whereas blocks and sticks were used in rib and bracket formation used to

strengthen the fuselage.


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WINGDESIGN AND CONFIGURATION

Aspect Ratio=b/c

Where b = wingspan, c = chord length.

The decision of the aspect ratio to be 6 was first based on the published experimental data.

The first estimation of the chord length was derived from the fact that the flight of the aircraft

would be made at an approximate Reynolds No. of 3, 00,000 and an assumed altitude of

3000 ft., standard day conditions and a flight speed of 51 ft. /sec, the = 0.002175 slugs/

ft3, = 0.3677x10-6 slugs/ft.-sec, then the approximated value for the chord length came

out to be 11 inches, which was deduced with the help of the airfoil data. The initial

wingspan was calculated with the help of the above information which further came out to

be 72 inches.

The solution obtained by solving the parametric equations for the optimum lift and drag

conditions, gave the chord length to be 12 inches and the wingspan to be 72 inches out of

which only 67 inches of wingspan is effective for generating lift and the rest 5 inches are

used in covering the fuselage which generates zero lift. The horizontal stabilizer provides lift

for the remaining surface area. Coefficients for lift, drag, and twisting moment of the wing

vary with the Reynolds No. and hence the angle of attack was found using airfoil design

software. These coefficients were verified with the published experimental data.

The structure of the wing should be designed such that it is sufficiently stiff to handle

highest loading impact encountered during takeoff and landing. The transfer of the wing

load should be span wise through a double beam bracket with the ribs enclosed in it and

chord wise through the ribs itself. The double beam bracket is the structure which has a

beam shaped like the leading edge part of the airfoil and the rear beam shaped like the


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trailing edge part of the airfoil. The double beam bracket configuration is used to counter

the twisting moments resulting from the generation of the lift.

The structure of the wing was chosen as shown above (the center part of the wing was kept

solid).

The solid part of the wing was sized 10.4 inches, extending to 5.2 inches on the each side

of the wing starting from the center of the wing. The dimension of the solid part of the wing

was determined by the conducting the stress analysis on the wing by finite wing method

encoded in the design & simulation software used i.e. Solidworks. For the wingtip to be

used in the aircraft the Hoerner wingtip was selected.


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Hoerner wing tip

The aspect ratio of the wing that the air measures is almost always less than what is

obtained when measured with a measuring tape. The difference is influenced considerably

by the wing tip i.e. if the wing tip is properly shaped than the difference would be small. By

employing the Hoerner wing tip in the design, the vortex cores which whirl off the tips are

kept as far apart as possible hence their placement is drifted to a point as far aft on the

chord as possible.

Although a high wing configuration is slightly less efficient than the mid wing configuration,

it adds to the stability of the aircraft, on the basis of which a high wing configuration was

chosen. A small dihedral angle was also employed in the design of the wing as it further

adds to the maneuverability of the aircraft during turns.


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FUSELAGE DESIGN AND STRUCTURE

The fuselage of the aircraft was designed keeping in view the performance and ease of

manufacturing. Design criteria required for the fuselage include ability to withstand the

different forces that it would encounter during the flight mode, take-off and at the time of

landing. The vital criteria included aircraft maximum gross weight, structure which should be

light yet strong, ease of manufacturing and operation. As the material for the fuselage

construction was chosen as Balsa wood (due to its characteristics), wood blocks of different

sizes and sheets of different dimensions were selected for building different parts of the

fuselage. The shape of the fuselage was chosen to be of box type configuration derived

from strength analysis of the beams constructed from the material chosen, it was deduced

that the balsa sheet of thickness 0.25 inches used as the box configuration is enough to

support the maximum load which could be encountered during any mode of operation. The

dimension of the fuselage was found out to be 49.4 inch long 5.5 inch deep 5.5 inch

wide.

The structure of the fuselage is supported by the bulk-heads and ribs which are strategically

placed inside the fuselage at the position where the nodes were obtained when finite

element method was conducted for the stress analysis.

Example of a Bulk-head or Rib


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The compartment where the payload is to be placed is enclosed within the walls of balsa

sheet of thickness 0.25 inches from the three sides and is then covered by the wing

structure from the above.

Maximum load enacting upon the fuselage will be supported by the Rib pieces or Bulk-

heads. Some of the Ribs were placed so as to bear payload whereas some were placed for

structural rigidity. The bulk-heads were made of the Balsa sheet of thickness 0.25 inches

since outer dimensions of the Ribs depended upon the dimension of the cross-section of

the fuselage where the Bulk-head was placed. The testing of the fuselage with bulk-heads

was done by applying 200 lb. and 80 lb. of compressive and tensile forces, respectively.

The above mentioned forces are considered sufficient keeping safe operation of the aircraft

in view and then whole fuselage structure was analyzed using Solidworks 2009. The results

of simulations clearly suggested that the box configuration along with the ribbed structure

was an effective and weight saving solution for the criteria defined. When the wing, cargo-

weight and landing gear are attached to the fuselage, the forces on the aircraft become

centralized. This centralization allows slightly stronger, heavier components to be used for

greater load distribution points.


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TAIL SELECTION

Tail comes into act as soon as the control of yaw and

pitch are taken into consideration. This is constructed

in the same way as the wing but in two distinct parts

for a better mounting of vertical stabilizer.

WHORTMAN FX 63-137 will be employed in making

the tail providing neutral lift (selection performed by

the same procession as that of the main wing) along with maintaining a well-disposed

lift/drag quantity. The size of the tail was calculated by taking care of the fact that the center

of gravity results to be closer to the front wheel and lying on the payload bay assembly,

necessarily to control the pitch in order to maintain an optimum angle of attack to produce

maximum lift. Whereas the drag should be kept to a minimum value i.e. the vertical

stabilizer assembly as well as the rudders should be kept as thin as possible and the area

should be kept very small so that the cross winds are not considerably affecting its yaw

movement. There are three types of tail planes used in a canard structure. The left most

one being cruciform type, the middle one being fuselage mounted and the one on the right

being tail mounted. Fuselage mounted being the strongest (as the drag related issues are

the least in slow speed high lift planes) was the most preferred by us.


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TAIL DESIGN

The three views shown here are vital for comparison of the dimensions. Left most diagram

shows the front view, the diagram in the middle shows the top view and the diagram at the

extreme left shows the side view.

This is the isometric viewof empennage


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ENGINE

The design vitae are incomplete without mentioning engine. Supporting it O.S. 0.61 FX has

itself been prescribed in the SAE rulebook because our planes fall under size 60 category.PARTICULARS OF ENGINE

2-STROKE OR 4-STROKE

2-STROKE engines are easier to use (maneuverability) whereas 4-STROKE engines are

more fuel efficient. The competition requires greater maneuverability.

RINGED OR ABC

Ringed Engines - An iron/aluminum piston moves inside iron sleeve, surrounded by rings

that provide compression.Advantages i) economical, ii) good start, iii) greater power.

Disadvantages i) Longer break in periods, ii) susceptible to damage on improper carburetoradjustment.

ABC Engines - An aluminum piston that moves inside chrome plated brass sleeve. The fit

of the piston and cylinder is perfected at the factory to provide excellent compression.

Advantages i)shorter break in ii)less susceptible to damage on improper carburetoradjustment

Disadvantages Costlier repair in case of damaged carburetor


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O.S. 0.61 FX PARTICULARS

The Bore length of 0.61 FX is 24mm and the Stroke length of 0.61 FX is 22mm which

calculates the RtSR ( ROD to STROKE Ratio) to be equal to 1.9091 which is appropriate

for a size 60 engine.Displacement 10 Cubic centimeters, 0.61 cubic inchesWeight 550 GramsType ABCStroke 2-StrokeRPM range 2,000 - 17,000

CAD Views of O.S 0.61 FX

The procession leads us to the very next step of designing the engine on CAD (computer

aided design) as a part of design responsibility. The four views have been presented below.


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CONTROL MECHANISMThis is the first stage encountered by every individual while proceeding for the designing of

an airplane, and is the foundation of the most important aspect when it comes to flight

controls.The prime control elements being the ailerons, rudder,

elevator and throttle. On the wing each aileron (used to

produce roll) has its own servo. These are assembled

and connected in such a way that they act opposing

each other.Rudder (used to produce yaw) is located atthe trailing edge of the vertical tail, the servo controlling

Rudder also controls the movement of the rear wheel

in order to allow steering on ground. Elevators (used to

produce pitch) being located on the trailing edge of the

horizontal part of the tail make use of one servo. The

throttle is located on the carburetor, controls the power

output of the engine by restraining the fuel supply to engine.


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SERVO SELECTION

The servo consists of a i) circuit board (miniature), ii) a combination of metal gears and

iii) a seemingly small but powerful electric motor. The horn(round disc above) is directly

linked with the gear assembly, this horn or the hand is connected directly to the control

surface of the plane by the means of a rigid servo linkage designated the rod.

Coreless Vs. Brushless

CORELESS: This design is lighter resulting in quicker acceleration and deceleration. The

result is smoother operation, and faster response time (for planes involved in acrobatics).

BRUSHLESS: This design is efficient, provide more power and speed. Offsets being

response time and smooth operation (for planes requiring reliability in hard weather

conditions).

For selection of servos the team employed this technique

Torque (oz.-in) = 8.5E-6 * {C^2. V^2. L. sin(S1) tan(S1) / tan(S2)}

Which further supported the usability of

FUTABA S3003 and FUTABA S3010 in the

aileron, rudder, elevator and carburetor throttle

control.


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LANDING GEAR

I.STRESS ANALYSIS

LandingGear undergoes maximum strenuous activities while landing so it must have great

strength. The basis behind the loading prediction for the landing gear is impulse momentum

formula:

(1)

So landing will be in series of square impulses as the plane hits and then bounces until it

rolls flat on the runway. There are three basic types of landing gear used in UAV out of

which tail gear is most suited for our fuselage as tail gear gives it stability as well as

structural safety. As we can see in the following representation of tail gear center of gravity

(C.G.) is b/w two front tyres and tail drag tyre as it facilitates the landing and reduces impact

forces.


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Wheel span of tail drag is much higher as it facilitates the landing gear by absorbing the

impact while landing. Displacement curve of center of gravity (C.G.) is much lesser in

magnitude in this case when compared to nose gear and bicycle type. After having

simulated different landing gear designs, we arrived at the conclusion that tail dragger

provides us with a larger range of impact angle thus making it safer for the pilot to land the

aircraft without compromising on its structural integrity.

After going through different variations of the tail dragger itself, the design meeting most of

our requirements was the conventional tail dragger but it had its own shortcomings which

led us to design an enhanced version on this landing gear by employing the concept of

cantilever in it. On performing analysis on cantilever type landing gear the vital prospect

which came into play was that the maximum stress bearability rose to three rimes as that of

the conventional type. Since this gear transforms shear stress into angular stress.

II. SELECTION OF MATERIALS

The two most important aspects of a landing gear are:

Material of landing gear selection

Material of tyre selection

Material required for the cantilever type gear should have an optimum composition of

tensile strength and shear stress fatigue point, so the two metals which outshine others

when it comes to these vitae was iron( specifically malleable iron) and steel( specifically


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annealed steel). The stress analysis has been shown below to further support our point of

view. It helped us land up to the result that annealed steel was a better option.

Selection of tyre was an equally important decision as the two parameters which are to be

taken into consideration are:

Area of contact( as greater area of contact would increase the rolling frictional force )

Tyre material( as the shock absorbing capability greatly depends on material )

Stress analysis of the conventional and cantilever tail dragger is shown below. And as we

can deduces from the data analysis cantilever had many advantages over it competitor .it

can withstand more stress with less deformation which will result in controlled landing with

less bounce.

Stress analysis of both the landing gear stimulated in solidworks is shown in isometric

views as follows:

Malleable Iron Annealed Steel


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Conventional Landing Gear

Cantilever Landing Gear


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DRAG ANALYSIS

We approximate the aircraft drag polar by the expression

CD = CDmin+ KCL2 + K (CL - CLmin)

2

The CDmin is made up of the pressure and skin friction drag from the fuselage, wing, tails,

landing gear, engine, etc. With the exception of the landing gear and engine, the C Dmin

contributions are primarily skin friction since we take deliberate design actions to minimize

separation pressure drag (i.e. fairings, tapered aft bodies, high fineness ratio bodies,

etc.).The second term in the CD equation is the inviscid drag-due-to-lift (or induced drag)

and K is the inviscid or induced factor = 1/( AR e). The e in the K factor can be

determined using inviscid vortex lattice codes such as AVL. The e for low speed, low sweep

wings is typically 0.9 0.95 (a function of the lift distribution).

The third term is the viscous drag-due-to-lift where K is the viscous factor = fn(LE radius,

t/c, camber) and CLmin is the CL for minimum wing drag. Both K and CLmin are determined

from airfoil data. The K term is difficult to estimate. It is usually determined from 2D airfoil

test data. The CDminterm is primarily skin friction and the data given in Nicolais White paper

will be used in its estimation.

The boundary layer can be one of three types: laminar, turbulent or separated. We

eliminate the separated BL (except in the case of stall) by careful design. For Re < 105 the

BL is most likely laminar. At a Re = 5x105 the BL is tending to transition to turbulent with a

marked increase in skin friction. By Re = 106 the BL is usually fully turbulent.


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ItemPlanform area

(in*in)

Wetted area

(in*in)

Reference

length (in)Cdmin

Fuselage 163 670 33 0.019

Engine 15 100 Na 0.004

Wing 864 1728 12 0.017

Horizontal tail 200 400 8 0.0007

Vertical tail 0 87.5 6.25 na

Landing gear 12 24 Na 0.0028

Total Cdmin =0.043.

Assuming a wing efficiency e = 0.95 gives an induced drag factor K = 1/( AR e) = 0.0335.

Notice that the often omitted viscous drag factor K = 0.0133 is 40% of the induced drag

factor. The total drag expression is

CD

= 0.043 + 0.0335CL

2 + 0.0133(CL 0.7)2


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PAYLOAD PREDICTED GRAPH


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DIMENSIONS AND STRUCTURAL VIEW


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TIMELINE & BUDGET


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REFERENCES

1. Society of Automotive Engineers. SAE Aero Design 2011 Rules and Guidelines.

2. Anderson, John D., Jr.: Fundamentals of Aerodynamics, 4th edition, McGraw-

Hill Book Company.

3. Woods, Graham. Foam Materials.

4. Woods, Graham. Wing Construction for Vacuum Bagging.

5. Simons, Martin. Model Aircraft Aerodynamics. Biddles Ltd, Guildford and

Kings Lynn.

6. Filippone , A.

7. Roskam, J. Airplane Flight Dynamics and Automatics Flight ControlsParts I

and II.

8. Crowe, J. H. Tandem-Wing Aeroplanes: An Examination of the Characteristics

of this Type of Wing Arrangement. Aircraft Engineering.

9. Budynas, Richard: Mechanical Engineering Design, 8th ed., McGraw-Hill Book

Company.

10. NASA. Ailerons.

11. Beer, Ferdinand: Mechanics of Materials, 4 ed., McGraw-Hill Book Company.

12. GraphiteStore.com.

13. Fieldman, Jim. "Great Planes Patty Wagstaff's Extra 300S ARF Product

Review.

14. Selig, Michael, Summary of Low Speed Airfoil Data, SoarTech Publishing.

15. Nicolai, Leland M. Estimating R/C Model Aerodynamics and Performance.
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Nsit Sky Knights - Design Report(048) (1)