Dust Thrusters

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Submitted November 16, 2007

Conceptual Design Report

Dust Thrusters

Dain Christensen - Cost/Components/Operation

Julene Forner - Structures

Jess Howe - Stability and Control

Jonathan Newhall - Configurations

David Roman - Performance

Mike Straka - AerodynamicsKyle Vonnahmen - Propulsion

AE 440


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Table of Contents:

1. Nomenclature

2. Executive Summary (JN)

3. Structures (JF, DC)

4. Propulsion (KV)

5. Aerodynamics (MS)

6. Stability and Control (JH)

7. Performance (DR)

8. Configurations (JN, DC)

9. Cost (DC)

10.Conclusion

11.References


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1. Nomenclature:

Adisk=disk area of propeller

Afcsw=fuselage cross sectional area at the wing aerodynamic center

Afcst=fuselage cross sectional area at the tail aerodynamic center

Av=area of the ultimate wakeb=wing spanb

h=span of horizontal tail

BHP=brake horse power

bv=height of vertical tail

c=wing mean aerodynamic chordch=horizontal tail mean aerodynamic chord

cht=horizontal tail area

CLL=derivative of the lift of horizontal tail wrt angle of attack

cm=derivative of pitching moment

cmfus=derivative of pitching moment of the fuselage with respect to

angle of attackcp=power coefficient

ct=thrust coefficient

cr=chord length at the root of the vertical tail

ct=chord length at the tip of the vertical tail

cvt=vertical tail volume ratio

Czamin=minimum total normal force

Czamax=maximum total normal force

D=propeller diameterFp=derivative of fin force wrt %alpha

fssw=fraction of the wing embedde in the slipstream

fT=corrects for non-zero thrust

g=gravitational constantGW=gross weight

J=advance ratioK=gust alleviation factor

K1=propeller downwash factor

K2=propeller downwash factor

KF=empirical pitching moment factorkp=propeller diameter coefficient

Lf=total length of the fuselage

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xp=position of propeller

zt=perpendicular distance between the tail aerodynamic center and the

line parallel to the x-axis and from the wing trailing edge

h =total downwash and derivative at the horizontal tail

p

=total downwashderivative at the propeller

CNblade

=normal force per blade with the propeller at zero thrust

=downwash angle derivative

p

=derivative of the downwash produced by the propeller slipstream

at the tailu

=upwash derivative at the propeller due to the wing

sa=efficency of shock absorber

h=efficiency of horizontal tail

p=prop efficiency

T=efficency of the tire as a shock

=wing tip to chord ratio=density

mass=airplane mass ratio

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2. Executive Summary (JN):

This design report is the first submission towards completion of the AIAA request for

proposal: Agricultural Unmanned Aircraft System (AUAS). The basics of the RFP [1] are to

create an unmanned aircraft and accompanying equipment that will spray a field with chemical

or solid particles. The aircraft must be capable of servicing a field which is 2640 by 1000 feet

and be able to take off and land on an unpaved surface which is a maximum of 750 feet long and

50 feet wide. The aircraft is to carry a payload of 300 pounds in addition to the equipment which

will apply the load. The final major requirement is that the aircraft and all equipment needed to

operate the vehicle must be moved from site to site with only a standard pick-up truck.

Instead of looking at three radically different designs in great depth, the group decided to

only carry the three designs into the first stage of development and then choose a configuration.

With this design, the next steps of the design process included extensive trade studies and

typically tried to look at three or more variations that could be applied to the situation. For

example, the wing for a traditional design is typically in a high, low, or mid fuselage

configuration. Each of these possibilities were considered and the best design for this set of

requirements was then chosen. This design philosophy allowed the group to continually narrow

and perfect the design instead of having the best features mixed with the worst features across

three designs. It also allowed for more in depth analysis since time was not wasted on

unnecessary calculations involved in three individual designs.

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3. Structures (JF, DC)

Introduction

The objective of this report is to establish a base for the future structural design of the

aircraft. A rough sizing for the fuselage was conducted along with a design for the structure of

the fuselage. The wing and wing box were selected based upon the ability to carry the critical

load paths. Materials were selected for the fuselage, wings, and skin of the aircraft. The landing

gear was selected based on trade studies. A V-n diagram was created using initial data for the

aircraft. Finally shear and bending moment analysis were conducted for the root of the wings.

3.1 Fuselage Sizing

To get a rough idea for the initial size of the fuselage, historical data from other aircrafts

were used. Usually the size of the fuselage is designed using the engine size, the cowling for the

engine, the payload, placement of the wings and tail, along with any cut outs needed in the

structure [2]. Without this data it was decided to use historical data to get a very rough estimate

of the fuselage size. Looking at the Piper Pawnee A-25-235 which has a length of fuselage to

length of wingspan ratio of .6878, and the Cessna 152 with a length of fuselage to length of

wingspan ratio of .7109, and the Piper Cub J-3 with a length of fuselage to length of wingspan

ratio of .6374. It was decided using the historical data and a wing span of 25 ft that the length of

the fuselage would be 16 ft. The rough height of the aircraft was decided in a similar fashion.

From the middle of the aircraft to the ground is 4 ft. The radius of the front of the fuselage was

designed to be 2 ft, using initial engine data that the engine would be roughly around 4 ft in

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diameter. This was a rough fuselage sizing, with more data on the sizing requirements of the

engine and payload, the actual dimensions of the fuselage will be established.

3.2 Fuselage StructureFour types of fuselage structures were considered when decided upon the configuration

for the fuselage structure. The structure types are the monocoque, the semi-monocoque, the

veneer, and the truss structure. When deciding upon the fuselage structure the following

considerations were taken into account: maintainability, strength, and its ease of production.

The monocoque type structure contains no longerons, struts or panels. Instead it is a

continuous shell usually made up of small wood strips laid on top of one another so the grain of

each layer is in the opposite directions. This type of structure is very time consuming to make

and requires more time to create than the other three configurations [3]. It is also difficult to

repair if the structure breaks because it is one continuous shell. Due to the time required to

produce the monocoque type structure and the maintenance problems it was decided to not use

the monocoque structure.

The next structure that was considered was the semi-monocoque type structure. This

structure is similar to the monocoque structure, but instead of being one continuous shell it has

two halves of a shell that are glued or bolted together. Like the monocoque type structure it takes

a lot of time to build, and is difficult to repair if it is broken. Therefore it was decided to not use

the semi-monocoque type structure.

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The next structure considered was the veneer structure. The veneer structure usually

contains four longerons, and uses elliptical bulkheads attached to the longerons by glue or

screws, for support. It also uses veneer panels attached to the longerons and the bulkheads for

added support [3]. Like the monocoque type structure maintenance would be difficult due to the

panels. The veneer structure also contains rings as its bulkheads which would be harder to create

or find than metal rods. Veneer is also not an ideal material in moister climates; the wood gets

moist and rots. This would not create a very suitable structure because the structure relies

partially on the panels for its strength. For maintenance difficulties and the fact that it uses veneer

panels it was decided to not use the veneer structure.

The final structure type to be looked at was the truss structure type. The truss structure

type usually contains four horizontal longerons held together by vertical struts. It uses triangles to

create a rigid structure. A benefit of the truss structure is it is easier to repair when damaged.

Unlike the monocoque and veneer there are not round bulkheads or rings which would be

difficult to replace. Also unlike the semi-monocoque there are not multiple stringers that would

be difficult to replace. There are only four longerons which if broken would be difficult to

replace, and smaller struts which are easy to fix if broken. The downside of using the truss type

structure is it is more box like than the others and therefore is not as aerodynamic. This can be

solved by using a small number stringers and small bulkheads to form the fuselage to be a

rounder shape. It was therefore decided that the best option for the structure of the aircraft would

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be the truss because of the high strength to weight ratio, and easy maintainability as seen in

figure 3.1.

Figure 3.1: Fuselage structure

The truss structure is designed to contain 45-45-90 degree triangles. This is so each

member carries the tension and compression loads equally, and no member is doing more work

than another. The vertical spars for structural stability should be spaced between 10 to 24 inches

apart. It was decided to go with 24 inches. It is also designed so that the longerons carry the

bending load, while the vertical and diagonal struts carry the shear. The diagonal struts are also

designed to carry the tension and compression loads, so that when one of the members is in

tension, its sister member is in compression.

3.3 Wing Structure

The structure of the wing needs to be designed to handle bending moments, tension,

compression, shear, and the load due to lift. The spars are designed to carry the bending load,

the web or covering of the wing handles the shear load. While the ribs are designed to maintain

the cross section and prevent the wing from buckling. The tension and compression loads are

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carried by both the spars and the ribs, while the load due to torsion is carried by wire bracing

Figure 3.2: Spar Design

between the ribs [4]. The front spar should be placed at a length of 15% of the chord length from

the leading edge, in our design .525 ft from the leading edge. The rear spar should be placed at a

length of 60-70% of the chord, in our design 2.275 ft from the leading edge [4]. This can be seen

in the below figure of the wing structure.

There two types of spar designs were looked at when choosing the type of spar for this

aircraft. They were the solid spar and the I beam spar. The solid spar tends to be heavier than the

I beam spar because it is more solid. However the solid spar is much stronger than the I beam

spar design. With the solid spar there is also the option of drilling holes through out the spar to

reduce the weight of the solid spar, this changes the strength of the solid spar by an insignificant

amount. It was therefore decided to use the solid spar as the spar design because of its strength

and ability be able to adapt to a lighter weight.

Another consideration for the structure of the wings is the placement of the ribs. The ribs

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should be place between 6 and 14 inches apart. This distance depends on the aircrafts

performance. The faster the aircraft flies the closer the ribs need to be. For our design, because of

the low speed, it is estimated to have the ribs spaced farther apart. This spacing reduces the

number of spars, which reduces the weight. Further structural analysis will be needed to get an

exact placement of the spars.

3.4 Attachment of the Wings

The RFP calls for an aircraft that can easily be transported in the back of a standard truck

or a in a trailer. Three different designs were considered to meet this requirement. These designs

include the foldable wing, a design where the tail and wings would be removed, and the design

were just the wings would be removed.

It was decided that the design where the tail and wings would be removed was also not

ideal for the design. This design is generally seen in small UAV designs and is not practical for a

larger UAV. This design would add weight to the fuselage because of the need to adapt the tail

end to have the tail be removed. Also, the fuselage is only estimated to be around 16 ft long. This

is long enough to fit in the back of a truck or trailer.

This leaves two designs for the wing attachment. It was decided for this report to have two

designs for the wing attachment. They are the foldable wing and the detachable wing. Both

currently proved to be equal when a rough trade study was conducted. It was therefore decided

that further information would be needed to decide upon the wing attachment. Both are designs

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The design also calls for a detachable part of the wing at the end of the wing, so it more readily

fits in the back of a truck. Both designs can be seen below figure 3.3.

Figure 3.3: Wing Designs

Further trade studies and analysis will have to be conducted to decide upon the design for

the attachment of the wings. Currently both are equally ideal.

3.5 Wing Box

A wing box is needed to account for the bending moment at the root of the wing, due to

the lift on the wing. There were four designs that were initially considered when picking a wing

box design. These designs include the box carrythrough, the ring frame, the bending beam and

the strut braced. When deciding upon the type of wing box, the following criteria were

considered, weight, space, ability to handle the sprayer system, and drag.

The first design considered was the ring frame. This design although it tends to reduce

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the amount of space the wing box takes up, requires heavy bulk heads that would substantially

add to the weight of the aircraft. It also is generally used for fighter aircraft designs and not

general aviation [5]. Due to the high weight of the ring frame wing box, it was decided another

design would be used.

The next wing box considered was the bending beam. This design relies on the fuselage

to carry the lifting load, and the beam to carry the bending moment. A benefit of this design is it

does not take up as much space as the box carrythrough design. This design is usually seen on

aircrafts composed mostly of composites due to the heavy strain the fuselage must endure due to

the lifting load it carries [5]. It was therefore decided another design would be considered.

The next design considered was the strut braced design. This design is usually seen on

general aviation, and is a great way to save space in the fuselage. However, it has a higher drag

than the other three designs because of the strut. This design will be used if the foldable wing

design is used for the wing attachment. It will save space in the fuselage, and reduce the load at

the root of the wing.

The next wing box design considered was the box carrythrough. This design is usually

seen on general aviation aircrafts. It is just an extension of the wing through the fuselage. This

can be seen in the below figure. This means a bending moment is not applied to the fuselage of

the aircraft. The wing box carries both the lifting load from the wings, and the bending moment

at the root of the wing. However the problem with the wing box carrythrough configuration it

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takes up valuable fuselage space. Compared to the other three designs this configuration best fits

the first fuselage structure that was selected. Therefore the box carrythrough design was selected

for the wing box design of the first wing attachment. This can be seen in the figure below.

Figure 3.4: Wing Attachment

3.6 Materials of Aircraft

Three different types of materials were considered when decided upon the materials to

make the aircraft fuselage out of. They include wood, aluminum, and steel. These materials were

chosen due to there relatively low cost compared to other materials available.

The first considered was wood. It has a good strength to weight ratio, and can act like a

composite due to the grain, for a much cheaper cost. Another benefit of wood is that structurally

it is easy to repair if damaged, and has a relatively low production cost [5]. It is generally the

structure used on homebuilt type aircrafts. One of the disadvantages from a structural point of

view is that in moister climates wood has a tendency to rot [5]. This rotting reduces the structural

integrity of the aircraft. Also if it gets wet the wood can expand due to moisture content which

deforms the structure, and can add weaknesses to the design. Our aircraft will be operated in

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third world countries where climates tend to be more tropical and moist. There for it was decided

that a wood frame would not be an ideal structural material for our aircraft.

This leaves two materials, steel and aluminum, to be considered for the main material of

the aircraft. Both materials are readily available throughout the world which would make them

ideal from the stand point of reparability of the material. Aluminum has a high strength to weight

ratio, and is relativity light [5]. Steel is heavier, but is stronger and has a higher fatigue resistance

than aluminum. It is also easier to fabricate steel, which means that steel comes at about 1/6 th the

cost of aluminum [5]. Steel tends to be seen on many truss design type structures, due to the fact

that it does not require as frequent maintenance as wood [2]. After completing a trade study

taking into account, strength, stiffness, density, corrosion tendencies, reparability, cost and

availability, it was decided that steel would be the ideal material for the fuselage. It is lower cost

and stronger than aluminum, it is easier to repair, and has a three times higher elasticity than

aluminum [5]. Another reason that steel was chosen and the initial design material for the

fuselage is it is much easier to convert from steel to aluminum than vice versa. Therefore during

later trade studies if it is decided that there is an ability to use aluminum, it will be easier to

convert to.

The next consideration is to look at the materials to create the structure of the wings out

of. For the same reasons as above it was decided to crate the spars of the wings out of steel. For

the ribs three different materials were considered wood, steel and aluminum. A composite rib

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taken into account, ability to handle on rough surfaces, prop clearance, and drag.

It was immediately decided that the multi-bogey configuration was not going to be used

as the design. There is no need for multiple tires, the aircraft is a lightweight aircraft. Also the

increase in tires would add increased drag. It was decided that the multi-bogey configuration was

to extreme for our design.

The quadricycle and bicycle configurations were also easily ruled out for possible

designs. Both require the aircraft to land and take off with a flat attitude [5]. Due to the

elementary controls that will control the aircraft this would be a difficult requirement to meet, so

it was decided another configuration could meet the needs of the design with out an increase in

the control system

The single main configuration is also not an ideal design for the aircraft. Its design has

the aircraft close to the ground. This design is not ideal for a crop duster because of the rocks,

dirt, and grass that can easily be thrown into the propeller if it is close to the ground [5]. The

single main configuration also is not ideal because it does not balance on its tires till the aircraft

starts moving, therefore when it is no in motion it rest on one of its wings. This would not be

ideal for an aircraft that needs to operation on rough terrain.

The landing gear configuration was therefore narrowed down to two configurations the

Tricycle configuration and the Taildragger configuration. A trade study was conducted to decide

on the type of configuration that would best suit the design. Taken into account was propeller

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clearance, ground handling, drag, weight, ability to handle rough terrain, takeoff distance, and

historical use on crop dusters. Some of the benefits of the Tricycle configuration were it can not

nose over due to the front wheel, and it is stable on the ground, it does not have problems with

ground looping [5]. Some of the negatives of the tricycle configuration were it has more drag due

to the front wheel and is heavier than the taildragger. Some of the benefits of the taildragger

configuration were it handles well on rough surfaces, it has a high propeller clearance, it has a

shorter takeoff distance than the Tricycle configuration, and less drag due to having a small rear

tire [5]. With the trade study conducted it was decided to go with the taildragger configuration, it

would best suit an aircraft that would need to operation on rough terrain, and take off in short

distances.

With a landing gear configuration chosen the type of shock absorbers was the next

decision. Six different types of shock absorbers where initially considered. The types include the

rigid axle, solid spring, levered bungee, Oleo shock-strut, Triangulated, and Trailing link

configurations. They were narrowed down to two configurations, taking into account cost,

maintaining, and ability to handle rough terrain.

The first designs to be ruled out for the possible shock absorber configuration were the

Triangulated, Trailing link, and Oleo shock-strut. The RFP calls for a low cost rugged vehicle

which is easy to maintain. The Oleo strut does not fit this criteria, it is higher maintenance, and

higher cost than the other three configurations. Therefore if the design used an Oleo strut the

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configurations was immediately discarded.

The next configuration to be discarded was the rigid axle. Although this design is low

maintenance and low cost, compared to the other two configurations. It did not compare when it

came to its ability to handle rough terrains. It relies solely on the tire for the shock absorber

which would not be enough shock absorption for our aircraft applications [5].

This left two configurations for the shock absorbers, the solid spring and the levered

bungee. A trade study was conducted taking into account the weight, drag, scrubbing effects, and

simplicity. The Solid Spring was the final choice because of its reduced drag, and its simplicity

which makes it easier to repair.

Next a rough estimate for the stroke, S, of the shock absorbers was calculated. It needed

to meet the requirements of an stroke between 8-12 in., this value is the general aviation value

from Raymer. Using

from Raymer and the fact that for general aviation vertical velocity is 10 ft/s, gear load factor is

3, efficiency is .5, tire efficiency is .47 and the calculated stroke of the tire to be .3354 ft, the

rough estimate for the stroke of the landing gear is .983 ft or 11.79 in. This fits well into the

required range of 8-12 in.

The final thing to be considered initially for the landing gear is the location of the landing

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(3.1)Sstroke=V

2vertical

2gNgear

T

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gear. The landing gear needs to form a 16 to 25 degree angle with the CG. Using this analysis, it

was decided that the landing gear would be placed 1.46 ft in front of the CG. For the taildragger

configuration the landing gear needs to be placed at a 25 degree angle from the front center of

the aircraft [5]. The final consideration is the tail wheel. According to Raymer the shimmy of

the tail wheel needs to be reduced so that when the aircraft is moving the tail wheel does not tear

off from the body of the plane. To solve this problem a rake angle of negative 4-6 degrees is

required.

3.8 V-n Diagram

When in flight the aircraft will experience basic loads, limit loads and gust loads. It is

important to know and understand the maximum loads that the aircraft can experience. A V-n

diagram shows the limit load factor as function of the airspeed of the aircraft [5]. A V-n diagram

is needed to ensure that the aircraft can structurally operate at certain velocities under certain

loads. The positive limit load factor and the negative limit load factor were calculated using the

following equations:

Using a gross weight of 1000 lbs, the positive limit load factor was found to be 4.28 and the

negative limit load factor was found to -1.71. Next using an Excel spreadsheet and the Cl, Cd, and

Cm-t data from aerodynamic data the maximum and minimum lift curves were calculated using

the following two equations given in class. The first being for maximum lift, and the second

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(3.2)

(3.3)

npos=2. 124 ,000GW10 ,000

nneg=.4npos


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being for minimum lift.

Where the maximum Cza is 1.41, minimum Cza is -0.664, density is 0.002378 slug/ft3, and W/S is

5.413. Next the dive speed was calculated using the fact that the dive speed is usually 40-50

percent higher than the cruise speed. Next the positive and negative gust load lines where

calculated using equation 3.6 as given in class:

Where KU is 17.5 ft/s, a is 5.167 1/radians, the velocity is 75.6 knots, and W/S is the same as

before. The gust loads and the maximum and minimum lift curves where then plotted to get the

following V-n diagram.

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(3.4)

(3.5)

n=Czamax

1

2W

S

V2

n=Czamin 1

2W

SV2

(3.6)ng=1

KUaV

498 WS


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Figure 3.5: V-N Diagram

The area enclosed by the blue curves is the region in which the aircraft can structurally

fly. Outside of this region the aircraft is no longer structurally sound and will break apart. The

green lines indicated the gust loads. The point on the V-n diagram labeled High AOA is the point

where the maximum load factor is reached without stalling at the slowest speed. This point is

important because here the load on the wing is almost perpendicular to the direction of flight.

The dive speed is also an important point because represents the maximum load factor on the

aircraft. At this point the aircraft has a low angle of attack which means the load is almost

vertical in the body axis. Both the High AOA and the dive speed can be used for structural sizing.

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It also can be noted that according to Raymer at lower speeds the load factor is limited by the

maximum lift that the aircraft can produce, at higher speeds the load factor is limited by the

positive limit load, based upon the aircrafts use (i.e. general aviation, transport, fighter).

3.9 Shear and Bending Moment Loads at Wing Root

The reaction force and moment at the root of the wing were calculated assuming static

equilibrium. The weight of both wings is approximately 167 lbs which means 83.5 lbs per wing.

The lift generated by the wings is 1000 lbs, and that implies 500 lbs per wing. The system on the

wing was represented as in figure 3.6, the weight of the wing is represented by a linearly

distributed load, and the lift was represented by a triangular load across the wing, with the lift

decreasing linearly, as it goes across the wing.

Figure 3.6:force and moment diagram for the wing

To simplify things the weight of the wing is considered a point force at the mid point of the wing.

The lift was represented as a point force, one third of the way down from root of the wing. These

modifications can be seen in figure 3.7.

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Figure 3.7 Simplified force and moment diagram for the wing

To find the reaction force and moment the following equations were used (the moment is solved

at the root)

Solving for these equations, yielded a reaction force of 416.5 lbs, and a reaction moment of

1311.63 lbs*in at the wing root.

Conclusion

Through this structural analysis many initial plans for the structural design of the aircraft

were decided. These include rough dimensions of the aircraft, the decision to use a truss structure

for the fuselage, steel to be used as the material of the fuselage and strut of the wing, aluminum

to be used as the rib of the wing, and fabric to be used as the covering. Two designs for the wing

attachment and wing box were decided upon. An initial V-n diagram was constructed to show the

maximum load the aircraft can handle at particular velocities. Finally bending moments and

shear were calculated at the root of the wing. Further analysis and trade studies will have to be

conducted on the structural designs of the aircraft, but the initial designs are now in place.

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(3.7)

(3.8)

0 500 83.3y Ff R= = 0 83.5 * 63 500 * 4 2y MM R= = +


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4. Propulsion (KV)

Introduction

The design requirement indicated that the aircraft needed to take off and land from a 750

ft by 50 ft. gravel or grass runway. In addition, the operating altitude for the aircraft was only 20

ft. AGL. Finally, the selected engine must be an off the shelf engine that can be easily acquired

[1].

4.1 Engine Selection

For basic aircraft, there are three engine options, a turbojet, turbofan, and propeller driven

engine. For this design turbojet and turbofan engines did not seem practical or cost effective. Due

to the low operating altitude and harsh take off and landing conditions, a jet engine would be

highly susceptible to damage caused by dirt and debris being sucked into the engine. So it was

decided that a propeller driven aircraft would be the best option. Once a propeller powered

engine was decided upon, the next thing was deciding whether to use a piston engine or a Wankel

(rotary) engine. A turboprop engine could have been used, but again it would not have been very

cost efficient.

4.1.1 Piston Engines

A piston engine is the basic combustion engine that has been widely use in aircraft since

the first flight of aircraft, and also used in automobiles. Piston engines used in aircraft today are

basically just modified automobile engines that are similar in design to engines used 50 years

ago. Due to the invention of the jet engine there has not been much research and development

that has gone into updating the piston engine for aircraft use. Piston engines however have been

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proven to be reliable engines. However, they need to be constantly adjusted during flight to insure

proper fuel and oil mixtures. The Piston engines also arent designed for the constantly high

RPMs that aircraft require, making them more likely to overheat or break. Therefore they also

need to be overhauled quite often, usually every 1200 hours or less. And since they are so

complex, it takes a long time and a lot of money to overhaul them [6].

4.1.2 Rotary/Wankel Engines

A rotary (Wankel) engine is a fairly new concept designed specifically for use in aircraft.

For example, rotary engines are designed to operate for long periods of time at high RPMs. This

is vital for aircraft. Another plus for rotary engines is that they have only a few moving parts,

compared to piston engines that have thousands of moving parts. This makes them less

susceptible to breaking. Rotary engines are also smaller in size and weight compared to piston

engines but provide more output power. They also cost less than piston engines and are cheaper

to maintain as well. The only disadvantage of rotary engines is that the combustion phase is

shorter than for a piston engine, resulting in a less efficient burn of the fuel which results in

slightly lower fuel efficiency. However current research is constantly improving the fuel

efficiency [7].

4.2 Design Choice/Methodology

From the initial sizing and constraint analysis it was estimated that the aircraft would

require a T/W ratio of around 0.3. For the analysis, it was difficult to accurately calculate the fuel

weight fractions due to the uncertainty of the fertilizer drop during the cruise part of the

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mission. Therefore we were only able to estimate the amount of fuel burned during this time. We

also had to take into account the possibility where the payload is not dropped and the entire

mission is carried out with a full payload. It was also calculated that the TOGW of the aircraft

was around 1000 lbf.

To calculate the required power needed for the aircraft, it should first be noted that the

term T/W (thrust-to-weight ratio) is used to describe jet engines while P/W (power-to-weight

ratio) is used to describe propeller engines. For homebuilt, singe engines, and/or agricultural

aircraft, a P/W of 0.07 to 0.09 is desired with units of (hp/lb) [8]. Therefore for an initial

estimate, the engine for the aircraft should be about 70-90 bhp to meet the initial design

requirements.

The next thing that needed to be calculated was the propeller diameter and the number of

propellers to be used. For our analysis, a prop system of 3 propellers was chosen to go along with

the data given in reference [8]. When comparing the number of propeller blades used, static

thrust and forward thrust will be affected. A two-bladed system will have a 3% better propeller

efficiency than a two-bladed system but with have about 5% less static thrust. A four-bladed

system though will have 5% better static thrust than the three-bladed system but a 3% reduction

in propeller efficiency for forward thrust [8].

To calculate the diameter length, several equations were used:

29

D=KP4Powerbhp ft (4.1)


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Table 4.1. KP Values (hp,ft units)

Number of Blades KP

2 1.7

3 1.6

4 1.5

Equation 4.1 and Table 4.1 were used to calculate the propeller diameter based on

horsepower and the number of blades used. Equations 4.2 and 4.3 are used to calculate the

propeller diameter at which the tips will go supersonic. The propeller diameter should be kept

under this value. For metal propellers V tip,helical needs to be less than 950 ft/s, less than 850 ft/s for

wood propellers, and less than 700 ft/s if noise is an issue. Equation 6.1 and Table 6.1 are also

used to calculate the propeller diameter based on engine output and number of blades. After

calculating the diameter using these two methods, the smaller diameter is then chosen as the

initial propeller diameter.

After calculating the diameter, the static and forward thrust can then be calculated. To

calculate the thrust, Figures 13.11 and 13.12 were used from [8]. To calculate the thrust the

following equations were also used:

30

Vtip, helical=Vtip, static2 V2= n D 2V2

Vtip, staic= n D ft/ s (4.2)

(4.3)


31/78

Equation 4.4 was used to calculate the advance ratio and equation 4.5 was used to

calculate the power coefficient. These two values were used to find P, the propeller efficiency,

and 3/4, the angle of the propeller of the diameter away from the hub, from Figure 13.12 in [8].

CP was then used to find CT/CP using Figure 13.11 from [8]. This data was calculated and

tabulated for the various engines being considered and for a different number of blades as well.

The engines selected for initial consideration were the Rotax 912 S piston powered engine, the

AR682R rotary engine, the Rotamax 650cc rotary engine, the Revolution One Turbo 650cc

rotary engine, and the Rotamax 1300cc rotary engine. Table 4.2 below shows the data and

variables calculated for each engine.

31

(4.4)J=V

n D

CP=550bhp

n3 D5

Tstatic=CTC

P

550bhp

nDlbf

Tforward flight=550bhpP

Vlbf

(4.5)

(4.6)

(4.7)


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Table 4.2 Engine Comparison Data

Rotax 912 S AR682R

Rotamax

650cc

Revolution

650cc

Rotamax

1300cc

Type Piston Rotary Rotary Rotary Rotary

Options w/EFI Turbocharger Turbocharger CarburetorHP 95 90 120 120 90 120

RPM 5500 7000 8000 6500 5500 6500

Weight (lbs) 164.6 124.6 130 140 170

Cost 18,000 11,000 13,725 11,070

SFC (lb/bhp/hr) 0.54 ~0.5 ~0.5

Fuel Flow(gal/hr) 6.8 3.9 - 7.6 4.5 - 5.5 2.8 - 5.6 4.92 - 9.84

Gear Ratio 2.43:1 2.03:1 2.273:1 (est.) 2.273:1 (est.) 2.273:1 (est.)

RPMprop 2263.4 3448.3 3940.9 2859.7 2419.7 2859.7

RPSprop 37.7 57.5 65.7 47.7 40.3 47.7D2 (ft) 5.31 5.24 5.63 5.63 5.24 5.63

D3,max 5.00 4.93 5.30 5.30 4.93 5.30

D4 4.68 4.62 4.96 4.96 4.62 4.96

Dmax,metal 7.98 5.24 4.58 6.32 7.47 6.32

Dselect 4.85 4.85 5.15 5.15 4.85 5.15

J 0.48 0.32 0.26 0.36 0.45 0.36

Cp 0.15 0.04 0.03 0.07 0.12 0.07

n,prop 0.55 0.55 0.50 0.55 0.60 0.58

Ct/Cp 1.15 2.50 2.75 2.15 1.75 2.00Tstatic,2 312.00 421.77 509.74 549.21 420.74 510.89

Tstatic,3 328.42 443.97 536.57 578.11 442.88 537.78

Tstatic,4 344.84 466.16 563.40 607.02 465.03 564.67

Tforward,2 336.36 318.66 386.25 424.88 347.63 444.19

Tforward,3 326.56 309.38 375.00 412.50 337.50 431.25

Tforward,4 316.77 300.09 363.75 400.13 327.38 418.31

4.3 Engine Results and Data

After comparing the different engines, the Rotamax 650cc rotary engine, was decided as

the engine to use. This engine was the cheapest of the five engines when comparing the output

power. It was also decided that a 3-blade propeller system would be the best. If 2 blades would

32


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have been chosen, the forward flight thrust would be higher but the static thrust would be lower.

The opposite would be true for a 4-blade system. Therefore a 3-blade system was selected to get

the best of both. The calculation of thrust for the selected engine revealed that the resultant thrust

was more than what was needed. When looking at smaller engines initially, their power output

was just barely enough to meet the requirement. So it was decided to ere on the safe side and

upgrade to a larger class of engines. Also, since the RFP it how there should be room for

improvement and development in the future. Therefore, if a customer decides to increase the

payload, they will not have to worry about getting a new engine in the process. Also for this

analysis, drag from the cowling was not taken into consideration. So the actual thrust that the

engine produces might be less than the calculated thrust.

Figure 4.1: Drawing and Model of the Rotamax Single Rotor 650cc Rotary Engine.

For the selected engine, the Rotamax Single Rotor 650cc Rotary Engine, shown above in

Figure 4.1, its turbocharged engine gives it an additional 10 lbs of boost. Figures 4.2 and 4.3

below show the difference between the engine having and not having a turbo charger. The

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turbocharger allows more air to enter the combustion chamber, increasing the energy released

from the combustion therefore producing more power. From the Figures 4.2 and 4.3 below, it can

be seen that a turbo charger has a big effect on the output power of the engine. Figure 4.4 below

shows the thrust versus velocity curve for the Rotamax 650cc engine. This graph shows that there

is a lot of available thrust at lower RPMs. This also shows that a smaller engine may be used. See

section 4.7 for more in depth discussion.

Figure 4.2: HP/Torque vs. RPM With Only a

Carburetor or EFI System

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Figure 4.3: HP/Torque vs. RPM with Turbocharger

Figure 4.4: Thrust vs Velocity Curve for the Rotamax 650cc Rotary Engine

4.4 Fuel System

This semester, not much time was able to be spent on the fuel system. For specific fuel

consumption, SFC, all that could be found was that the engine had an estimate SFC of 0.5. This

35

Thrust vs. Velocity

0

200

400

600

800

1000

0 10 20 30 40 50 60 70 80 90 100

Velocity, V (ft/s)

Thrust

,T

(lbf

Actual Thrust

Cruise Speed

Static Thrust

Theoretical Thrust


36/78

was provided by the Rotamax Company. However from the engine data it was found that the

selected engine had an estimated fuel burn rate of 5.5 gallons per hour. It was estimated that the

mission would take about one and a half hours to complete. With a 20 minute fuel reserve and

noting that not all of the fuel will be able to be extracted, it was decided to calculate the fuel

needs for a two hour mission. Since rotary engines are capable of using regular 87 octane

gasoline, an estimated total of 11 gallons of fuel would be needed for the mission. Using the fact

that 7.5 gallons of fuel occupies 1 ft3, the aircraft will require 1.47 ft3 of fuel space [8]. Also using

the fact that one gallon of gasoline weighs approximately 6 pounds, the total weight of our fuel

will be about 66 pounds.

When considering where to put the fuel in the aircraft, it the group chose between putting

it in the wings, in the fuselage, or a combination of both. Since the design calls for the wings

detaching for easier transportation, it was considered to be easier to just put the fuel in the

fuselage. Putting the fuel in detachable wings would require an extensive system of hoses and

connectors that would be prone to leaks after repeated use in the rugged landing conditions.

However, putting the fuel in the fuselage would result in a lot of empty space in the wing. What

was decided upon was to extend the wings out approximately 3 to 4 feet and then have the wings

detach after that point. This would allow fuel to be put inside the wings, using a discrete tank,

without having to worry about fuel connectors. Also if there is still leftover fuel in the wings,

they can be detached leaving the leftover fuel in the wings. If the fuel were in the detachable

36


37/78


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5. Aerodynamics (MS)

Introduction

The basic requirements for the aerodynamics of the aircraft were taken from the AIAA

2007/2008 RFP. The design calls for a very simple design that can fit in a confined space. The

purpose of the aircraft is to deliver liquid or solid material onto a field in order to spray

chemicals or drop fertilizer on the crops. Two primary requirements for the design forces are the

ability to fit in a trailer and also that the design and construction must be easy to maintain and

operate. These requirements led to the following analysis of different wing geometries, with the

results following.

5.2 Wing Geometry

The wing geometry was selected with simplicity and cost in mind. The aerodynamic

efficiency of a given geometry was taken into account, but not as significantly as the prior

requirements. First, the sweep of the wing was taken into account. A wing that is swept back

generally has a lower drag for equal lift generated. But a swept back wing also introduces a more

complex design and would cost more to acquire and also be more complex to maintain. As a

result, the wing would have no sweep of the leading edge. Another geometry consideration is the

taper ratio. By the same arguments as for the sweep, a taper ratio of 1 was decided for the wing.

The use of an unswept and untapered wing also enables simpler internal structures to support the

weight of the fuselage while in flight.

The placement of the wing was also considered. A high-wing (mounted to the top of the

38


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fuselage) versus low-wing (mounted on the bottom of the fuselage) were both considered. The

deciding factor between these two designs was dictated by the purpose of the aircraft. One of the

two main purposes of the aircraft is to spray chemicals on an agricultural field. Booms to carry

and spray the liquids need to be carried over the entire span of the wings and even further out to

increase the spray area over each pass, so the low-wing design was selected in order to

accommodate the booms. Attaching the booms to the wings will keep weight down, since the

booms do not need a structure to support their own weight.

A wing-tip is also an important aspect of a wing. The three designs that were considered

are: rounded, sharp, and cut-off tips. According to Raymer, a rounded tip would easily allow air

to flow around the tip, which will decrease the efficiency due to increased induced drag. A sharp

edge would make it more difficult for the air to flow past the tip, but would require caution when

working around the ends of the wing. The sharp design is also another component that would

have to be considered when looking at maintenance. Hence, the cut-off wing-tip was selected for

the unmanned crop-duster. The cut-off wing tip provides better resistance to airflow around the

tip and also offers the simplest design out of the three already simple designs [9].

In order to calculate the size of the wing, initial sizing data was used to perform the

calculations. From initial sizing, the wing loading of the aircraft is 11 lbs/ft 2. As a result, the

wing reference area is 90.9 ft2. The span of the wings was limited to 25 feet, and as a result the

aspect ratio of 7 was selected. This aspect ratio falls in between the values that Raymer provides

39


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based on historical data [9]. The historical data suggests an aspect ratio of 7.5 for agricultural

purposes and 6 for homebuilt/ultralight, both of which are appropriate descriptions for the

unmanned crop-duster design. The chord length ended up being 3.6 feet as a result. A drawing of

the wing with the dimensions is provided in the configurations section.

5.3 Airfoil Selection

When the airfoil selection came under consideration, an airfoil that has lift over a wide

range of angle of attacks was desirable. At spraying speed, the angle of attack for the aircraft

must be stable enough not to cause loss of control. Since the aircraft is flying only 20 feet above

ground, any stalls might lead to a crash, so a high angle of attach at which the aircraft stalls

would be desired in order to prevent a crash. The database provided by Selig [10] was used in

deciding on a proper airfoil that would provide the performance that is demanded of an

unmanned aircraft of this type.

The airfoil that was selected is the S4083; this airfoil is shown in figure 5.1 [10]. The

airfoil has a good overall lifting force, while maintaining relatively low drag through a wide

range of alpha. The CL versus angle of attack for this airfoil can be seen in figure 5.2 with a

Reynolds number of 2.10 x106, with a Mach value of 0.01. The values were obtained using

XFOIL. The maximum angle of attack that this airfoil allows before stall is more than 13 degrees.

This will provide the grounded pilot more maneuverability in which to operate the aircraft

without going into a stall.

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Fig. 5.1 Airfoil coordinates of S4083

Fig. 5.2 Wing Lift Curve of the S4083 Airfoil

When considering the performance of the airfoil at the take-off and landing, stall speed

was important to calculate. According to Raymer, the stall speed can be calculated by using the

general lift equation, but using CLmax rather than CL.

41

(5.1)


42/78

The maximum lift coefficient is known, so calculating the stall speed becomes trivial. The stall

speed becomes 24.4 fps. This speed is very reasonable for the application of an unmanned crop-

duster, since the landing speed will be 31.8 fps. This is a relatively low speed and will enable the

ground pilot enough time to make maneuvers to land the aircraft safely. With this analysis, it was

decided that high-lift devices would not have to be employed in order to provide better stall

speeds when landing/taking off. Also, the wing incidence angle to the fuselage will be set to

zero, since the lift at an angle of attack of zero is high enough to support the cruise speed of the

aircraft.

5.5 Parasite Drag Computation

The parasite drag from the component build-up method used is given by Roskam [11].

This method takes the sum of the individual components and refers it to the aircraft via the

reference area of the wing. The basic formula for finding the CDo is given as:

In this equation, CD is given in table 5.1 and the A as the reference area to use for the

calculation. The CD values are given in Roskam and apply to propeller driven airplanes. Since

this applies only to the surfaces of the aircraft, the landing gear term will be added.

42

(5.2)


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Table 5.1 Parasite drag buildup method

Component CDo A (given in ft2)

Wing .0070 90.9 (Sref)

Fuselage .1100 8 (total fuselage frontal area)

Tail, horizontal .0080 14.8 (Sh)Tail, vertical .0080 8.28 (Sv)

Interference Add 5% to CDo

Roughness and

Protuberances

Add 10% to CDo

The total calculated parabolic drag is then CDo= 0.0187, for the surfaces of the aircraft. In

addition, the landing gear will contribute significantly. The nonretractable landing gear of the

solid spring type is given a Cd value of 0.62 [12]. This value is corrected to the sizing of the

aircraft by multiplying it by the frontal area of the wheel and dividing by the reference area of the

wing. The landing gear drag coefficient ends up being 0.008. This value is added to the total

parabolic drag, so the total parabolic drag is now 0.0267. And an additional 15% is added to this

value to account for the spraying booms, roughness, and interferences, to give the final value of

0.031.

5.6 Drag Polar

The drag polar was created using the cambered formula given by Raymer. The Oswald

efficiency factor (e0) of .85 [1] is used to calculate the value of K.

The drag polar can be seen in Fig. 5.3 and demonstrates that the overall drag on the plane is

lowest when CL is equal to 0.7. Even though the design speed of the aircraft while flying over the

43

(5.3)

(5.4)


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field is higher, the drag is still very low overall.

Fig. 5.3 Drag Polar of the S4083

5.6 Lift-Curve Slope

The lift coefficient was computed using XFOIL software. In order to calculate the C l for

the airfoil, the slope was taken during the linear portion of the Cl versus angle of attack figure.

This value was then corrected using the following formula.

This formula was modified, since it also contained a term for the sweep angle, but would end up

being zero. Using the suggested value for (Sexposed/Sreference)(F) of .98, given by Raymer, the CL is

calculated and is equal to 4.49. This value also matches the sample values provided by Raymer.

44

(5.5)

(5.6)

(5.7)


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5.7 Future Work

Although the current geometry and airfoil selection satisfy the requirements, methods that

are more precise will be employed to maximize the performance of the aircraft, while staying

easy to maintain and simple to fly. More complex methods for predicting the accurate behavior of

the wings will be employed in the future to better predict the performance of the aircraft. The

addition of a boundary layer and other airflow properties will lead to more complex results and a

re-evaluation of the selected airfoil will ensue. Also, more accurate methods for predicting the

total drag of the aircraft will be evaluated.

6. Stability and Control

Introduction

Longitudinal stability of the aircraft is dependent upon many parameters. The initial task

was to determine the size of the control surfaces, such as the vertical and horizontal tail. Then,

the neutral point was found. A weight build-up was used to determine the center of gravity using

a simple weight times moment arm analysis. These values were used to ensure the static margin

was at an acceptable value for all flight conditions.

6.1 Tail Sizing

Historically, the vertical tail volume ratio is 0.04 and the horizontal tail volume ratio is

0.5 [13]. These volume ratios are defined in equations 6.1 and 6.2.

45

(6.1)

(6.2)

cht=

lh Sh

c S

cvt=lv Sv

bS


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Since different references suggest different methods to determine lh and lv, the average of

both methods was used [13]. The two methods suggest either using 2.5-3.5 times the wings

mean aerodynamic chord or 40-50% of the wingspan. Both methods yielded approximately the

same range for the tail length. The aspect ratio of the horizontal tail is typically 4-5 while the

aspect ratio of the vertical tail is typically 2.5-3 [13]. The aspect ratios for the tail are defined in

equations 6.3 and 6.4.

The rudder size was determined using the fact that the rudder area to vertical tail area

ranges ratio from 20-35%. The elevator size was determined using the fact that the elevator to

horizontal tail area ratio ranges from 20-35% [13]. The rudder size was used to determine the

root chord and tip chord of the vertical tail. The resulting geometry based on these equations and

given geometry of the wing are summarized in Table 6.1. When a range of values was given, the

average of these values was used.

46

(6.3)

(6.4)

ARh=

bh2

Sh

ARv=1.55

bv2

Sv


47/78

Table 6.1: Horizontal and Vertical Tail Geometry

l (ft) 11.07591

Sv (ft2) 8.282105

Sh (ft2) 14.77408

bv (ft) 3.833283bh (ft) 8.153734

cv (ft) 2.160578

ch (ft) 1.811941

Se (ft2) 4.062873

Sr (ft2) 2.277579

ct (ft) 0.594159

cr (ft) 3.726996

6.2 Determination of Neutral Point

In order to have static pitch stability, any change in angle of attack must generate

moments that oppose the change. This means that Cm must be negative. The point where the c.g.

can be placed and there is no change in pitching moment as angle of attack is varied is called the

neutral point. This is the aft most location the c.g. can be located to still maintain static pitch

stability. Equation 6.5 shows the definition of Cm used [14]. All positions are defined from the

tip of the fuselage with the positive defined in the aft direction.

Cm=CL XcgXacw Cm fushSh

SCLh

h

Xach Xcg Fp

qS

p

XcgXp

Moving the c.g. to the neutral point (Cm=0) and solving for the position yields equation

6.6.

47

(6.5)

(6.6)Xnp=

CL XacwCm fush Sh

SCLh

h

XachF

P

qSP

XP

CLhShS

CLhh

FPqS

p


48/78

Where CL, CL h , Xacw and S were provided from aerodynamic data. Sh was determined

above and q was determined from flight conditions. Cm fus was determined using equation 6.7.

In this equation, KF is an empirical pitching moment factor determined from Figure 16.14

of [14]. The maximum width of the fuselage, WF, the total length of the fuselage, LF, the mean

aerodynamic chord, c, and the reference area, S, were all given parameters from the geometry of

the aircraft. The efficiency of the horizontal tail, h, was different for power off and power on

conditions and is defined in equations 6.8 and 6.9, respectively.

The fsst factor is included to account for the area of the horizontal tail in the slipstream of

the propeller and is defined in the following equations 6.10 to 6.14. [13].

The total downwash derivative at the propeller is defined in terms of the upwash

48

(6.7)Cm fus=KFWF

2

LFcS

(6.8)

(6.9)

hT=00.9

h=

hT=01fsst TqAdisk

(6.10)

(6.11)

(6.12)

(6.13)

(6.14)

VdiskAdisk=VuAu

w=1

2 [VV2 2TAdisk ]Vdisk=Vw

Vu=V2w

fsst=1

bh 4 AuAfcst


49/78

derivative at the propeller due to the wing, which can be found using Figure 16.11 of [14]. The

total downwash derivative is then defined by equation 6.15.

The total downwash derivative at the horizontal tail is different for power off and power

on conditions and is defined in equations 6.16 and 6.17 respectively.

In this case,

was found using Figure 16.12 of [14] and is based on the aspect ratio, r,

m, and . The parameters r and m were found using equations 6.18 and 6.19, respectively.

The derivative of the downwash produced by the propeller slipstream at the tail was

found using equation 6.20.

NB is the number of blades. K1 and K2 are propeller downwash factors and were found

using Figure 16.17 in [14]. The propeller normal force coefficient were found using Figure 16.15

49

(6.15)

p

=1

u

(6.16)

(6.17)

h

=1

h =1

fsstP

(6.18)

(6.19)

r=lt

b

2

m=Z

t

b

2

(6.20)

P =K1K2NB

CNblade

P


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in [14]. The propeller normal force coefficient was also used to calculate the propeller normal

force, as in equation 6.21.

The function f(T) was found in Figure 16.16 of [14]. Once all of the parameters were

defined, the neutral point was found using equation 6.21. The neutral point was found to be 4.29ft

from the front of the fuselage with the power off and 4.27ft from the front of the fuselage with

the power on.

6.3 Determination of center of gravity

The location of the c.g. for each flight condition throughout the entire mission profile was

determined using the weight build-up from Configurations. The moment arm for each

component in each flight condition was computed using the angle of attack information from

Performance, where the front of the fuselage was the datum. The results are summarized in table

6.2.

Table 6.2: Center of gravity location for each flight condition

Flight Condition c.g. location [ft]

Takeoff 3.852694105

Climb 3.812109688

Cruise (beginning) 3.912126104

Cruise (end) 4.241919987

Descent 4.133202527Landing 4.241919987

50

(6.21)FP=qNB AdiskC

Nblade

f T


51/78

6.4 Determination of Static Margin

In order to have a neutrally stable aircraft, the c.g. must be located in front of the neutral

point. This means the static margin (as defined in equation 6.22) must be positive.

The static margin was initially negative for the cruise condition. In order to adjust for this,

the payload was moved closer to the engine. Since the fertilizer is flammable, a firewall was

added to the engine configuration. This added more weight to the front of the plane and the c.g.

shifted forward, resulting in static pitch stability for the cruise condition. The static margin for

the landing condition was also initially negative. The avionics box and the agricultural systems

sphere were shifted forward in the fuselage. The results for the static margin for the flight

conditions are summarized in table 6.3.

Table 6.3: Static Margin for each flight condition

Mission Segment

Static Margin

[%]

(power off)

Static Margin

[%]

(power on)

Takeoff 12.2 11.6

Climb 13.3 12.6

Cruise - beginning 10.5 9.8

Cruise - end 1.4 .68

Descent 4.4 3.7

Landing 1.4 .86

A trade study with a T-tail configuration was completed in order to see the effects of

increasing the perpendicular distance between the wing and horizontal tail. The results are

summarized in table 6.4.

51

(6.22)S .M.= Xnp Xcg


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Table 6.4: Static Margin for each flight condition with t-tail configuration

Mission Segment

Static Margin [%]

(power off)

Static Margin [%]

(power on)

Takeoff 13.7 13.1

Climb 14.9 14.2

Cruise - beginning 12.1 11.6

Cruise - end 2.9 2.5

Descent 5.9 5.5

Landing 2.9 2.5

The T-tail configuration proved to have a larger static margin than the traditional

configuration. The T-tail configuration is more likely to pitch up at high angles of attack when

the propeller is on than the traditional configuration, but this particular aircraft will not generally

be operating at high angles of attack. Further investigation into the cost and benefits of the T-tail

configuration is required.

6.5 Future Work

The next step for stability and control is to ensure lateral-directional stability. In addition

to that, the flexibility of the longitudinal fuselage, wing span-wise bending, and wing torsional

deflection must be considered. The dynamic stability will be the largest challenge for this

aircraft, as much of the mission profile includes turns. There is also a constant payload drop

throughout the cruise condition. This will provide additional dynamic instabilities. Ensuring

dynamic stability will probably involve the development of a stability augmentation system.

7. PerformanceIntroduction

Conventional crop dusters are often lower-end small aircraft models that an airplane

manufacturer may have on their product line. This is because simple designs are often suitable for

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the relative short range and low service ceilings of the missions flown during crop dusting. For

these reasons the RFP asks for an affordable, uncomplicated Agricultural Unmanned Aircraft

System (AUAS) capable of propagating liquid chemical fertilizer or solid seeds (or fertilizer

particles) to a plot of about 61 acres. In addition, the fuel reserves are to be designed to last a

total of 20 minutes of flight. The aircraft will cruise at an altitude of 20 feet to apply the

fertilizers to the field; however ferry flights require the aircraft to be airworthy at 1,000 feet.

Flight at this altitude will be discussed in the ferry mission profile section. The landing and

takeoff distances are at a maximum of 750 feet.

It can be inferred from the RFP that the performance requirements necessary to fulfill the

customers criteria are low when compared to general aviation. Consequently, this crop duster is

primarily designed around cost-effectiveness, simplicity in design, and reliability under varying

circumstances. These are the motivators for developing the performance analysis in the mission

profile segments.

7.1 Crop Dusting Mission Profile

Crop dusting requires a mission profile that is highly symmetrical. However, it differs

from other aircraft mission profiles in that the payload is not detached at one point during the

cruise segment but in this scenario the payload is dispersed throughout the entire duration of the

cruise segment. Figure 7.1 illustrates such a mission. Segment 1 is the warm-up and taxi five-

minute requirement before the 2nd segment (take-off). The aircraft must climb, segment 3, to its

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50-foot above ground level (AGL) requirement. Once it is aligned with the field the crop duster

will descend to a 20-foot AGL, segment 4, where it will spray the field until the entire area has

been covered in the 5

th

segment. Finally, segments 6-8 are climb, descent, and landing

respectively.

Figure 7.1: Mission Profile

7.2 Warm-up and Taxi

Segment 1 assumes the aircraft is built from the configuration in which it was transported

and is powered up. The performance requirements for this segment are accounted by the

propulsion system to be able to taxi and the landing gear to be able to move in different runway

conditions. These issues are discussed in their respective sections.

7.3 Take-off

The AUAS is not heavily regulated by the Federal Aviation Regulations (FAR)

airworthiness standards because of the low service ceilings, flight in urban areas, and unmanned

flight system. Nevertheless, a 50-foot requirement for clearing an obstacle was given by the RFP

to ensure the aircraft will clear possible telephone poles and two-story buildings in the

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surrounding area. Eq. 7.1 and its integral were used to numerically compute the take-off velocity

to be 108.9 feet/second and take-off distance to be 532 feet. In this analysis no intermediate

safety speeds were calculated because the aircraft is equipped with a single propeller. Refer to

Figure 7.2 for the computed values.

Figure 7.2: Take-off and Climb Lengths

7.4 Climb and Descent

There are two climb and descent portions during the mission profile. In the first portion,

the aircraft climbs from sea level to 50 feet then it descends to 20 feet. In the second portion, the

aircraft climbs from 20 feet to 50 feet then it descends to sea level. The velocity, given a rate of

climb dh/dt, is calculated via Eq. 7.2. It was calculated that a 13-degree climb and descent angle

for take-off and landing, respectively, was favorable. The average velocity during the climb

segments was 185 feet/second while it was 95 feet/second for descent segments.

55

(7.1)v t=v t 1

m[TDWL]t

(7.2)dv

dt=v t [g

T D

W1

v

dh

dt]t


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7.5 Cruise and Turns

Potential customers will most likely base the performance of this aircraft on its ability to

dispense fertilizers onto the field. Hence the fifth segment of the mission profile, as illustrated in

Figure 7.1, is the most critical in achieving product differentiation from other crop dusters on the

market.

Length and width requirements for a rectangle 61 acre plot of land are give to be 2640

feet and 1,000 feet respectively. It is estimated that at an altitude of 20 feet and a wingspan of 25

feet the sprayer system will have a swath width of 30 feet. Figure 7.3 demonstrates the proposed

mission path around the field. Since the swath width will have to cover the 1,000 feet width of

the field, the aircraft will have to make 34 passes. It is desirable for the aircraft to finish the entire

job (including turns) in 70 minutes in order to optimize the maximum flight time to be 1.5 hours

including the 20-minute fuel reserve. Given this constraint the aircraft can spray the field at 65

miles per hour during each pass. This value is well above the stall velocity.

A steady-level turn analysis was made using Eq. 7.3. The load factor n was capped by the

structural limitations of the aircraft. See the structures discussion. This allowed for up to 175

degrees/second for turn rate at a 76.4 degree bank angle. The mission requires two regimes of

turns. From the improvised runway to the field the sharpest turn radius is 20 feet while from one

pass to another the sharpest turn radius is 15 feet.

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(7.3)=g

vn21


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Figure 7.3: Mission Path

7.6 Landing

The discussion for landing is analogous to the discussion for take-off. Minor exceptions

occur in the landing flare complications. The descent and landing distances in Figure 7.4 account

for the flare, increased coefficient of friction and decreased velocity during landing. The short

landing distance is augmented by the reverse thrust provided by the power plant.

Figure 7.4: Landing Lengths

7.7 Ferry Mission Profile

This AUAS must be capable to fly ferry flights for special transportation and

miscellaneous missions. The mission profile is then modified to figure 7.5. Warm-up, taxi, take-

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off, and landing requirements remain the same for this mission profile. However, the climb,

descent, and cruise segments must be altered to account for a reduction in atmospheric density

from 2.37 x 10

-3

lb/ft

2

at sea-level to 2.30 x 10

-3

lb/ft

2

at 1,000 ft. This crop duster will not have

any design limitations to fly at that altitude. Usually in aircraft the limiting factor at higher

altitudes is the powerplant. In spite of this the drop in density is insignificant and the rotary

engine can operate under those altitude effects and also achieve the two-mile range required.

Figure 7.5: Ferry Mission Profile per RFP

7.8 Conclusion

This section attempts to quantify the aircrafts performance as it applies to achieving the

requirements of the customer. This AUAS meets all the requirements given by the RFP while

maintaining costs low. However, due to limitations in this report the discussion here is

rudimentary and details in calculations were omitted. As the project develops and more data

becomes available, an in-depth discussion into mission optimization will be presented.

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8. Configurations (JN, DC)

Introduction

The purpose of the configurations is to coordinate all of the other areas and bring the

design together into one unified airplane. The first step was to pick a design path and then to

narrow the design so that more advanced analysis was possible. The group looked at three more

radical designs against the standard configuration as the starting point for the project.

8.1 Initial Sizing

The initial sizing for the UAV was done using Aircraft Design: A Conceptual Approach

Forth Edition by Daniel Raymer [18]. The sizing was preformed to get an initial takeoff weight

(Wo) estimate for an aircraft that would meet the design requirements. The method in Raymer

starts by breaking the weight down in to four different components; The weight of the crew,

payload, fuel and the weight of the empty aircraft.

The weight of the fuel and the weight of the empty aircraft are directly related to the

takeoff weight, and as such the equation written above can be expressed as

The weight of the crew and payload were set forth in the RFP. Since this is an unmanned

aircraft, the crew weight is zero. The payload weight was set to 300 pounds, as per the

requirements set fourth in the RFP. It should also be noted that there was no consideration for the

payload drop during this analysis. This was so that if there was a problem with the delivery

59

(8.1)

(8.2)Wo=WcrewWpayloadWf

Wo

We

Wo

Wo=WcrewWpayloadWfuelWempty


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system the aircraft would still be able to land safely, also this would provide a lager margin of

error on the heavy side for the aircraft.

The weight of the fuel was found by performing calculations to find the fuel fraction

Wf/Wo (the weight of the fuel divided by the takeoff weight). First the mission profile had to be

analyzed, and broken in to segments, as previously discussed in section [18]. The mission has

three distinct parts; the takeoff, the crop dusting, and finally the landing. As per the RFP

specifications, no credit was given to the climb or decent. The crop dusting was represented in

this analysis as several consecutive cruise segments. The number of cruise segments required was

determined by taking the width of the field (1000 ft) and dividing that by the wingspan of the

airplane. According to Raymers method the fuel fraction for takeoff is .970. Raymer also lists

the Landing fuel fraction as .985. The cruse segment fuel fractions was calculated using the

Range of the cruse (.5 miles), the velocity of the aircraft (estimated to be 65 miles per hour, or 93

feet per second), the L/D (estimated to be 20), and the specific fuel consumption (estimated to be

(2.34E-5). The fuel fraction was determined to be .999, with 134 passes required to completely

spray the field. The fuel fractions were multiplied together to get the mission fuel fraction. The

RFP also called for 20 minutes of reserve fuel. This yielded a total fuel fraction Wf/Wo .067.

Table 8.1 has a short summery of the Mission fuel break down.

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Table 8.1: Fuel Fraction

Mission segment wi/wi-1

Takeoff 0.97

Landing 0.995

Each pass of the field 0.999

Wi/Wo for take off and first pass, no payload drop 0.97

Wi/Wo for take off and all passes, no payload drop 0.97

Wi/Wo for take off, all passes and landing, no payload drop 0.96

Wf/Wo (including reserve fuel) 0.067

The next step was to calculate the empty fuel fraction We/Wo This was done using in

iterative process. The gross takeoff weight equation above can also be expressed as

The weight of the Crew, Payload, and the fuel fraction are all ready known, so a Wo was guessed.

From this guess and historical data a We/Wo, can be calculated. This calculation was then

plugged in to the above equation and compared with the guess. Microsoft excel was used to do

the iterations. If the Guessed Wo and the calculated Wo did not match, the Guessed and the

calculated weights were averaged and the calculation was repeated using this average until the

two numbers were with in a pound of each other. The We/Wo was found to be .603.

From these numbers the initial takeoff weight was estimated to be 910 pounds. Since this

was just an estimation it was decided to ere on the side of caution and round the take off weight

to 1000 pounds for all initial calculations. Upon doing a weight build up later this number ended

up being a fairly good estimation of the aircrafts weight. table 8.2 has a breakdown of the weight

for each part of the aircraft.

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(8.3)W

o

=WcrewWpayload

1WfW

o

WeW

o


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Table 8.2: Finial weight estimates in Pounds

Crew 0.00

Payload 300.00

Fuel 61.25

Empty 549.14

Total 910.39

8.2 Design Comparison

Three different radical designs were considered from the outset of the project. These

were: a VTOL craft, a flying wing, and a twin boom pusher. Sketches of these designs are shown

in table 8.3. Basic research was conducted by the team to evaluate the effectiveness of each

design. To get a conceptual idea of how these ideas might work in the real world, different

categories were chosen to provide a basis for comparison. With the traditional design as the

baseline, numerical values were assigned to each quality so that a quantitative approach might be

taken. The results of this analysis are found in table 8.3. After careful consideration and much

discussion, the group decided that a radical design was not the best way to meet the demands of

the RFP. All three of the radical aircraft designs were given a negative rank, indicating that while

each had advantages for this application, overall they were not well suited to give the highest

performance across a broad range of requirements.

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Table 8.3: Design Comparison

Traditional

VTOL Flying Wing Twin boom pusher

Overall Cost-4- high composite

,makeup highpropeller costs

-2- composites arenecessary forcomplex structure

-N No componentsgreatly increase ordecrease overall cost

Safety

+1- propellers are in,cased in a ring

minimal risk ofinjury

+1- controlledlanding with poweroff

-N Generally is assafe as traditionalwhen no passengers

.are on board

Stability-N no problems were

found with stability

-3- requirescomputer softwareand control laws

N Cg is farther back

but not an issue

Maintenance-N no positive or

negative in this area

+1- large wettedarea allows easyaccess formaintenance

-1- Engine cooling is aproblem causingstress on enginecomponents

Durability+1- designed for the

,military can withstand a lot

-N no majoradvantages ordisadvantages

-1- Rocks and debrisare a problem in theprop

TOGW-1- as of now toolight for the

applications needed

+1- entire surfacelifts giving less

weight

-N Extra boomweight is minimal

Handling+1- ,handles well easy to maneuver

-N computercontrols allowease of hangling

-N wings lesseffective and tailmore effective due toprop wash

Sprayintegration

+1- already has hadspray cansintegrated in otherdesigns

+1- lots of area toinstall sprayer

-1- Propwash is wherespray systemoperates

Ability to fit

in truck

+1- wings are

,detachable smallenough to fit

-1- blended wingcan not be

separated fortransport

-N No major

advantages ordisadvantages

Drag-N No major

advantages ordisadvantages

+1- ,Minimal entirebody lifts

-N No majoradvantages ordisadvantages

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Total 0 -1 -3

8.3 Constraint Analysis

To get initial values of the thrust to weight ratio and the wing loading, a constraint

analysis was done. This analysis utilized equations in Raymer [19] as well as historical data for

similar class aircraft. The results of this analysis are best summarized in figure 8.1 below. From

this graph a design point was chosen which gave the group a rough estimation of the thrust

required for this vehicle as well as the most basic dimensions of the wing. This point is shown

below in figure 8.1 and has a corresponds to a T/Wo value of .3 and a Wo/S of 11. The constraint

analysis made many assumptions and estimations so that a basic design could be molded into a

more fully developed model.

0.0000 10.0000 20.0000 30.0000 40.0000 50.0000 60.0000

0.0000

0.2000

0.4000

0.6000

0.8000

1.0000

1.2000

1.4000

Wing loading vs. Thrust to weight

Cruise

Landing

Instant Turn

Take off

Sustained Turn

Wo/S

T/Wo

Figure 8.1: Design point analysis

8.4 Weight Build-up

With initial data in hand, the team set out to do more detailed analysis on the airframe. An

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airfoil was chosen and from that many other qualities were locked into place. An initial tail

sizing, following The Principles of Design [18], gave the final necessary to construct a detailed

weight build up. The weight buildup not only gave a more accurate estimate of the gross take off

weight but also broke it out into components so that the CG of the craft could be found. The

results of this analysis are found in table 8.4 below.

Table 8.4: Initial Weight Build Up

Wing 167 lbs. Engine 150 lbs.

Horizontal Tail 9 lbs. Fuel System 6 lbs.

Vertical Tail 13 lbs. Other 30 lbs.

Fuselage 70 lbs. Control System 40 lbs.

Main Landing Gear 52 lbs. Payload 500 lbs.

Tail Landing Gear 3 lbs. TOTAL 1040 lbs.

The more accurate estimate of 1040 lbs. is very similar to the original estimate of 1000

lbs. From this point the weight of the plane was broken down into components so that more

detailed analysis could be performed on.

8.5 Modeling

Once the basic structure of the airframe was decided upon, sketching and modeling

became necessary. Using the software package, Autodesk Inventor each of the components were

modeled and then joined in and assembly file. This visualization was important because it gave

the group a clear idea of what it looked like as well as provided an excellent means of

communicating the idea to others. The first view of the plane was only the most basic

dimensions. The first model is shown in figure 8.2 below

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Figure 8.2: First Visualization

The second iteration of the design incorporated more elements of the design as well as

changed the aesthetic of the plane by adding fillets. The front of the fuselage was modified to

include a rounded, lower drag front section and lead to a more aerodynamic shape as is seen in

figure 8.3.

Figure 8.3: Second Visualization

The third model in this design string includes the final modifications for the airframe.

Main landing gear were added to the underside of the fuselage. Also, the vertical tail was

redesigned; a tapered leading edge was added for decreased drag. The final model included as an

isometric view in figure 8.4 as well as a detailed (dimensioned) view on the next page.

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8.6 Conclusion

All of the internal components will need to be modeled in the next stage of the project.

Additionally, it will be necessary to modify the current design. As more and more detail is added

to the analysis, the design is sure to morph and change.

9. Cost Analysis (DC)

Introduction

The cost of any tool, compared to its expected value and usefulness is one of the major

considerations in weather or not it gets purchased. To this end it is important to find the expected

flyaway and operating costs of the proposed UAV. With this information, the consumers can

make an informed decision on weather or not this UAV will help them with there farming

operation.

9.1 Avionics and Operation

The plane will be operated from the ground using a standard joystick and keyboard. There

are already remote control airplanes that have software that allow the operator to practice flying

from their computer, using the actual remote control used to operate the aircraft. This UAV will

take it one step further and use the same software used to simulate flight on the computer to

control the aircraft.

Feedback will be provided to the pilot through two cameras, one mounted to each wing,

and a GPS located mounted in the fuselage. It is assumed the pilot will be the farmer who owns

the land, and that they are intimately acquainted with their fields. The GPS locator will show

where the aircraft is on a map, as well as show the UAVs velocity. A third camera will be

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mounted in the aft part of the fuselage so that the farmer can watch the spray fall and ensure

proper fertilizer distribution. The cameras will cost $300 each, and the GPS locater will cost

$800.

The information will be relayed back to the ground station via an antenna. A suitable

antenna can be purchased for under $200, as a very powerful antenna is not needed. This leads to

a total avionics cost of $1700. The software and ground station will cost about the same as a

modern personal computer, which is $2000. This brings the total cost for avionics and ground

operation to $3700.

Table 9.1: Avionics cost

Component Cost Quantity Total Cost

Camera $300.00 3 $900.00

GPS system $800.00 1 $600.00

Antenna $200.00 1 $200.00

Ground Station $2,000.00 1 $2,000.00

Total

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