Conceptual Design Report

Crop Dusting Unmanned Aerial Vehicle (AUAS)

Flying Unmanned Bug Annihilation Remote Vehicle

Team Members: Area of Responsibility: Adam Carrington Cost Analysis Glen Fetsch Stability and Control Eric Johnson Aerodynamics Brian Lin (Team Leader) Configuration/AutoCAD Phil Martorana Structures Therese Prosecky Propulsion Raj Ramachandramoorthy Performance

Table of Contents: Page # Nomenclature 3 1. Executive Summary 4 2. Configuration 5 3. Propulsion 12 4. Aerodynamics 21 5. Performance 30 6. Stability and Control 39 7. Structures 47 8. Cost Analysis 60 9. Weight and Balances 69 10. Conclusion 74 11. References 77 12. Appendix A 79

2

Nomenclature ng gust load factor a speed of sound or normal

coefficient curve slope Ngear gear load factor ηp propeller efficiency AGL above ground level P/Wo power-to-weight ratio AR aspect ratio RFP request for proposal AUAS ag. unmanned aircraft system Re Reynolds number b wingspan ρ∞ free-stream density bvt span of vertical tail

MR manufacturing cost per hour c chord length AC cost of avionics QR quality control cost per hour

Cbhp propeller sfc TR tooling cost per hour CDi induced drag coefficient

Sht Horizontal tail reference area CDL&P drag from loss and perturbance SFC specific fuel consumption CDmisc miscellaneous drag coefficient Sflapped “flapped” wing area CDo parasite drag coefficient Sref wing reference area ce/c elevator size STOL short takeoff/landing

EC cost of engine Svt Reference area of vert. tail Cf skin friction drag coefficient Swet wetted area FTC cost of flight tests Q interference factor

Q number of UAV’s to be built CG center of gravity Cl,max max lift coefficient T thrust Cl,max,l max lift coefficient at landing Tstatic static thrust Cl airfoil lift coefficient t/c thickness to chord CL wing lift coefficient TOGW takeoff gross weight Clα airfoil lift curve slope T/Wo thrust-to-weight ratio CLα wing lift curve slope U gust speed

MC manufacturing material cost UAV unmanned aerial vehicle V maximum velocity (knots) Cp power coefficient

cr/c rudder size Vc cruise velocity Cs speed-power coefficient Vd dive velocity

Vstall stall speed SSC cost of spray system Vtip speed at tip of propeller blade Ct thrust coefficient Vtip, helical helical tip speed of d diameter Vtransverse horizontal speed of propeller eo Oswald’s efficiency factor W weight of aircraft DP design point eW empty weight (lbs) FC fuel consumption

FF form factor Wengine weight of engine FTA number of flight tests W/S wing loading

We/Wo empty weight ratio MH hours of manufacturing Wf/Wo fuel weight ratio QH hours of quality control

TH hours of tooling

J advance ratio K gust alleviation factor lt tail length L/D lift to drag ratio Mtip Mach number of blade tip Mtransverse Mach of moving aircraft n+/- maximum positive or negative

load factor n rotation speed

3

1. Executive Summary (AC)

Crop dusters of the past were regularly used to spray fields with herbicide,

pesticide, and fertilizer; however with the development of more efficient ground based

sprayers crop dusters have all but extinct. These ground based sprayers are typically

more consistent, cheaper, and do not require the inconvenience of an airport or a pilots

license. These are the reasons why the use of crop dusters has become so rare.

The agricultural unmanned aircraft system (AUAS) overcomes all of these issues.

It is a versatile unmanned aerial vehicle (UAV) capable of fulfilling the farmers’ dusting

and fertilizing needs efficiently and affordably. It would initially be employed remotely

by an operator with minimal training. In future development, it may even carry an

autonomous autopilot system that would allow the aircraft to virtually fly itself on the

most efficient path possible to complete the mission. The lack of a cockpit reduces the

size and weight of the UAV allowing it to take off virtually anywhere.

This project’s goal is to develop an AUAS that is affordable for both private

farmers and developing countries. Additionally, it is desired that the aircraft will be

practical and easy to operate. The AUAS must be a fixed-wing, unmanned aircraft

capable of carrying an expendable payload that consists of 100 liters of chemical or

300lbs of solid particles. The aircraft must have fuel reserves for at least 20 minutes of

flight and have an operating altitude of 20 feet. The maximum landing and takeoff

distance is 750 feet on a 50 foot wide gravel or grass airstrip. It should be able to obtain

short flights of 1 to 2 miles with no payload at an altitude of 1000 feet. The aircraft and

any additional equipment should be able to be carried or towed by a pickup truck.

4

2. Configuration (B. L.)

2.1 Initial Configuration Selection

Initially, there were six proposed configurations which are illustrated in Figure

2.1.1. The configurations and reasons why they were nominated are as follows:

Flying Wing: Tow drag design for better performance and fuel efficiency. Initially the

flying wing was thought to have a wider wing span due to the lack of tail.

Bi-plane: The bi-plane was proposed because it is the stereotypical and therefore most

commonly known configuration for crop-dusters. This led group members to

believe that it was the most reliable configuration.

Conventional: The conventional configuration was chosen to be a baseline for

performance comparison. The conventional configuration’s simplicity in design

gives it its inert stability opposed to more radical designs.

Twin-Boom Pusher: The twin-boom design provides more structural rigidity because of

its enclosing box structure. The twin-boom also allows the storage of payload in

the two booms, separate from the center body. This provides more freedom of CG

placement for the airplane.

Conventional Pusher: The conventional pusher configuration was proposed because of

lower form drag and increased maneuverability. The pusher configuration also

allows the engine to be place closer to the CG allowing more leeway to place

other items in more various locations.

Swept-Wing-Conventional: The swept-wing-conventional configuration has lower drag

than straight wing configurations and also provides more stability and is capable

of higher speeds.

5

Flying wing

Bi-plane

Conventional

Twin-Boom

Pusher

Swept Wing Figure 2.1.1 Configuration Elimination and Selection

6

Three of these proposed configurations were discarded and three were kept. The

flying wing was disregarded mainly due to the requirement that all payload equipment

required that the equipment fit in a 1 ft sphere. This sphere would disrupt the

aerodynamics of the wing and therefore defeated the purpose.

The bi-plane configuration was not chose because two wings would produce two

sets of vortices which would induce large amounts of drag. Since our main goal is

performance and also affordability, the short coming of the tried and true bi-plane

outweighed its benefits.

The swept wing was not chosen simply because the purpose of the swept wing is

to allow the aircraft to achieve high seeds. The mission of the crop duster did not require

high velocity performance or maneuvers.

2.2 Initial Sizing

The initial sizing model was constructed from a simplified mission profile using

historical data for weight fractions for warm-up, take-off, climb, and landing. The crop

dusting leg of the mission profile was considered as a continuous cruise that would be the

length of cruise required to cover the whole field given a spray width equal to the wing

span of the airplane and an alternating pattern that repeats lengthwise along the field.

Table 2.2.1 Initial Sizing Input Data

Configuration Parameter

Name Twin-BoomConventional

Tractor Conventional PusherCruise Length (n.mi.) 16.44 16.44 16.44

Cbhp 0.4 0.4 0.4 Cruise Velocity (knots) 48 48 48

L/D 10.773 11.67 11.59 Reserve Fuel Requirement 0.05 0.05 0.05

7

Table 2.2.2 Historical Fuel Fractions Warmup+Take off 0.97

Climb 0.985Climb 0.985Land 0.995

Table 2.2.3 Initial Sizing Calculations for Twin Boom

Leg Discription R (n.mi.) R (ft) C (lb/s) V (knots) L/D Wi/Wi-

1 Wi/Wo1 Warmup+Take off 0.97 0.97 2 Climb 0.985 0.9555 3 Cruise 16.44 99891.34 2E-05 48 10.773 0.9977 0.9532 4 Climb 0.985 0.9389 5 Land 0.995 0.9342

Table 2.2.4 Initial Sizing for Conventional Tractor

Leg Discription R (n.mi.) R (ft) C (lb/s) V (knots) L/D Wi/Wi-

1 Wi/Wo1 Warmup+Take off 0.97 0.97 2 Climb 0.985 0.9555 3 Cruise 16.44 99891.34 2E-05 48 11.67 0.9978 0.9534 4 Climb 0.985 0.9391 5 Land 0.995 0.9344

Table 2.2.5 Initial Sizing for Conventional Pusher

Leg Discription R (n.mi.) R (ft) C (lb/s) V (knots) L/D Wi/Wi-

1 Wi/Wo1 Warmup+Take off 0.97 0.97 2 Climb 0.985 0.9555 3 Cruise 16.44 99891.34 2E-05 48 11.59 0.9978 0.9534 4 Climb 0.985 0.9391 5 Land 0.995 0.9344

Table 2.2.6 Calculated Weights and Weight fractions

Configuration Calculated Value Twin Boom Conventional Pusher Conventional TractorTOGW (Wo) (lb) 1470 1163 1163 Empty Weight (lb) 1155 848 848

We/Wo 0.786 0.729 0.729 Wf/Wo 0.069 0.054 0.054

Mission Fuel Weight (lb) 101.527 62.743 62.743

The mathematical equations used for calculating the initial-sizing are listed in refs [1].

8

2.3 Constraint Analysis

Figure 2.3.1 represents the preliminary initial constraint analysis using

preliminary and estimated input values. The initial design point was chose as one of the

lowest T/W0 values and the corresponding W/S0 value or 9 and 0.35 respectively. Using

the initial TOGW determined from the initial sizing and the chosen design point, a

preliminary thrust and wing area were computed. These numbers were then passed to the

other team members, namely propulsion and aerodynamics so they could begin their

initial calculations and estimation of values. Once the initial design point and design

weight were distributed to all group members, each group member produced revised

versions of the values assumed for the initial constraint analysis as can bee seen in Table

2.3.1. This produced a more refined constraint analysis as seen in Figure 2.3.2.

Initial Constraint Analysis

0

0.2

0.4

0.6

0.8

1

1.2

0 5 10 15 20 25 30 35

W / S 0

clmaxl = 3Take OffCruiseDP

Figure 2.3.1 Initial Constraint Analysis

9

Constraint Analysis

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0 5 10 15 20 25 30

W/S0

T/W

0

Take Off

Cruise

Landing

DP

Figure 2.3.2 Constraint Analysis and Design Point Table 2.3.1 Initial and Refined Constraint Analysis Input Values

Initial Once

Refined C l,max,l 3 2.342592 C l,max 1.775148 1.6268

CD0 0.797048 0.022 AR 8 8

Once the aerodynamics member received a wing area, three high-lift low-speed

airfoils were chosen: The Epler 431, Selig 2091, and Clark Y.Cl values were calculated

and then used for further refinement of the constraint analysis. The airfoils and their Cl

values can be seen in Table 2.3.2 below.

Table 2.3.2 Selected Airfoils and Cl values

C l,max,l C l,max e431 2.342592 1.6268

Selig 2091 2.31696 1.609 Clark Y 2.18592 1.518

10

An initial trade study was done to show the effect of using the different airfoils on

the constraint analysis. Figure 2.3.3 illustrates this effect. Apparently the three airfoils are

indeed very similar and have minimal effect on the landing constraint. This trade study

also illustrates that the magnitude of the Cl,max is proportional to the design space.

Landing Constraint Trade Study

0

0.2

0.4

0.6

0.8

1

1.2

0 5 10 15 20 25

W/S0

T/W

0

Landing e431Landing Selig 2091Landing Clark YSeries4Take OffCruiseDP

Figure 2.3.3 Landing Constraint Trade Study All equations used to calculate the constraint values and diagrams are listed in reference [1]. All airfoil geometries selected from references [2] 2.4 Future Work To further improve the design point and constraint analysis, the process of refining the

constraint analysis and optimizing the design space should be iterated several times in

order to converge to the best design point.

11

3. PROPULSION (T.P.) 3.1 Requirements from the Request for Proposal According to the guidelines set by the AIAA Request for Proposal (RFP), the

propulsion system parameters were not strictly specified. The only constraints on

designating a system were that the engine and its components had to be capable of

managing the requisite cruise and take-off speeds, as well as the gross weight of the

aircraft and payload. The engine and its operational systems had to be powerful enough

to carry at most three hundred pounds of seed and fertilizer and compensate for the

weight aerial vehicle.

3.2 Preferences for Propulsion Systems Initially, before any decisions were made, research was done regarding current

engine types being used in both conventional agricultural aircraft and in unmanned aerial

vehicles.

3.2.1 Radial Engines

It was discovered that conventional crop dusters used either radial engines or a

turboprop power plant. However, these engines, while capable of hauling the desired

weight, were too cumbersome for such an aircraft as the RFP called for. They had the

necessary horsepower, but due to their added output, were also about as heavy as the

vehicle that was to be designed. Added to this fact, they were able to output enough

power to generate speeds in excess of what was required and were therefore deemed both

uneconomical and impracticable.

12

3.2.2 Wankel Engines

Following this decision to discard the radial engine as a possible propulsion

system, the Wankel rotary engine was chosen as a candidate to be used in one of the three

aircraft designs. Some of the advantages of using a Wankel configuration were that this

type of engine had a smaller weight and produced more power at higher revolutions per

minute, as opposed to the radial engines. The Wankel rotary engine was also simpler in

design as well. It did not contain any valves or a crankshaft, but uniquely had a rotor to

allow for an increase in thermal expansion within the combustion chamber. It was

discovered that the Wankel did actually have some drawbacks in that the fuel injection

system was highly complex and the exhaust system released a large quantity of

hydrocarbons because of inefficiencies present in the engine. [3]

3.2.3 Unmanned Aerial Vehicle Engines

Upon completion of research concerning conventional aerial application aircraft

engines, studies were conducted into engines that dealt solely with light-weight,

unmanned aerial vehicles. Though the rotary engine did have some disadvantages, it was

ultimately the final choice. Two manufacturers, UAV Engines and ROTAX, were

investigated and engines were chosen based upon maximum horsepower generated, fuel

consumption, and weight. Shown in Table 3.2.3-1 is the list of engines considered during

the design process that would have been viable for the specific usage of agricultural

means. While the UAV Engines did have a high power-to-weight ratio and specific fuel

consumption, they did not have the power requirement to fly the aircraft. Eventually, it

was the ROTAX engines that were decided upon based on their produced power and

reasonable fuel consumption rate.

13

Manufacturer Model Weight (lb) Power (hp) Max. RPM Fuel Consumption (gal/hr)

UAV Engines AR801 53.7 50 8000 2.5 AR801R 65 51 8000 2.4

ROTAX 582 UL-2V 79.2 53.6 6000 3.302150638 582 UL-2V 79.2 65 6500 3.566322689 912 UL 121.3 79 5500 2.905892561

Table 3.2.3-1 Engine Considerations

Between the two UAV Engines engines, both running at about 50 to 51 horsepower, and

the three ROTAX engines, it was determined that the ROTAX rotaries would be best

suited to fit the requirements and constraints that were imposed upon the aircraft. Among

the choices were a 53.6 horsepower (hp), a 65 hp, and a 79 hp. The 79 hp engine was

equipped to handle the take-off and cruise power demands, but weighed approximately

121.3 lb. This weight was not suitable for the purposes of the design point and instantly

established the choice to lie with either the 53.6 hp or the 65 hp.

The first engine, the 53.6 hp, while lighter than the others, was deemed not

suitable for the allotted weight of the aircraft. Perhaps in future work, this engine will be

considered, but for now, the 65 hp is the prime selection for the propulsion system.[4]

3.3 Propulsion System Selection It was the ROTAX 582 UL-2V 65 horsepower rotary engine, as described in

Table 3.3.2, that was chosen to be the sole power plant for the agricultural aircraft.

Instead of using two lower horsepower engines, analysis proved that one engine with a

higher horsepower would be appropriate for use.

Manufacturer Model Length (in) Weight (lb) Power (hp) Max. RPM Max. Torque

(ft.lb) FC

(gal/hr) ROTAX 582 UL-2V 30.3465 79.2 65 6800 55.3 7.2

Table 3.3.2 Selected Engine

14

The 582 UL was a two-cylinder, two-stroke rotary engine that was temperature-

maintained via a self-contained liquid cooling system. This particular engine came

equipped with a Ducati dual capacitor discharge unit (DCDI) ignition system that, though

used primarily for motorcycles, was well-placed for use in the individual design

specifications. The fuel system was such that the aircraft could either use leaded or

unleaded motor octane (MON) 83 rated fuels. Upon investigation of this specific grade of

fuel, it was determined that premium unleaded fuel, such as that distributed from a local

gas station, would also be feasible as a fuel source. Also included of note, is an intake

silencer to limit the amount of noise generated by the engine and thus produce a more

aesthetically pleasing and efficient agricultural craft. [4]

3.4 Propeller Options It was indicated through researching that a propeller made of metal was far more

efficient than one composed of wood. With a wooden propeller selection, the efficiency

decreased by 10% for each blade. The propeller was chosen to be of constant-pitch

design with a pitch angle of at most 12° according to Fig. 13.12 in Ref [1]. Though the

inefficiency decreased for constant-pitch, it was decided that the issue of greater weight

for the constant-speed propeller farther outweighed its higher efficiency over the

constant-pitch. [1]

3. 4.1 Tip Speed and Mach Number Once the engine data was obtained and the type of propeller was chosen, the

integration between the engine and propeller had to be performed. The tip speed of the

propeller blade, as well as the maximum diameter, of the entire propeller, were first to be

15

calculated. The tip speed was crucial because it was then used to find the Mach number at

the tip of the blade. By rule, the Mach number at the blade tip cannot be greater than one.

To find the tip speed, while keeping the spherical tip speed at 950 feet per second (fps) as

suggested by research materials, the following equation was used:

2,tip tip spherical transverseV V V= +

2 (Equation 3.4.1-1)

However, with the larger propeller diameter and power supplied from the engine,

the blade tip Mach number was far greater than one and unacceptable. To account for this

result, a gear reduction system is usually installed along with these types of engines to

decrease the Mach number and keep it below one. With the addition of this reduction

package, there were some losses to the system that must be allotted for in any following

calculations. [1]

3.4.2 Propeller Geometry and Material With a proposed diameter of 62 in., a three-bladed configuration, and a 5 in.

projection for the hub, the weight was estimated to be about 8.03 lb. As mentioned

before, the propeller was designed to be made up of thin aluminum to coincide with the

rest of the body of the aircraft. Though most propellers are either wood or composite

material, it was inferred that wood had a lower overall efficiency and composite material

was far too costly. [1]

16

3.5 Performance of Engine Selection

3.5.1 Required Power With the design point chosen, the engine had to be able to accommodate a thrust-

to-weight ratio of 0.35. Coupled with the weight of the aircraft being approximately 1100

lb with payload, this placed the maximum thrust needed to be about 385 lbf. Figure

the engine. The fuel consumption of the 582 UL 65 hp is about 3.566 gal/hr based upon

information from the manufacturer and is only minutely higher than the 53.6 hp engine,

adding to the final decision to use the 582 UL.

3.5.1-1 shows the thrust versus velocity curve relevant to the total consumable power of

[4]

3.5.2 Power Ratios and Loading

0

50

100

150

200

250

300

350

400

450

0 100 200 300 400 500 600 700 800 900 1000

Veloci ty (kts)Figure 3.5.1-1 Thrust vs. Velocity curve

The maximum available power of the 582 UL-2V rotary engine is 65 hp, with a

power-to weight ratio of approximately 0.821 hp/lb. While the specific engine power-to-

weight ratio was slightly less than one, the total power to weight ratio of the aircraft was

far less than unity with a value of 0.059 hp/lb. This ratio, often interchanged with thrust-

to-weight ratio for jet engines, is semantically the same quantity only applied in terms

17

18

0

50

100

150

200

250

0 10 20 30 40 50 60 70

Power (hp)

Pow

er L

oadi

ng (l

b/hp

)

endent on the total weight of the aircraft, is graphically

t, the power loading ratio was at its maximum. This

roportionality of the ratio, which is lb/hp.

.6 Fuel System Integration

onfigurations, fuel systems had to be designed to fit space

Figure 3.5.2-1 Power loading vs. Power Supplied

befitting a propeller aircraft—an appropriate P/W value for an agricultural single engine

aircraft is about 0.08 hp/lb. [1]

The power loading, dep

shown in Fig 3.5.2-1 as a function of a varying power supply which was determined by

the nature of the mission profile.

At its lowest power outpu

result was expected because of the inverse p

The general desire was to stay somewhere in middle so as to achieve a required amount

of power to keep the aircraft in flight, but at the same time not cause the engine to out.[1]

3

For each of the three c

available aboard the craft, but also be placed in such a location that they did not disrupt

the center of gravity and become detrimental to the stability of the aircraft.

Both the conventional and pusher models had fuel tanks in the wings of the aerial

vehicle. There were three cylinders, all of the same dimensions, holding roughly

0.833197 US gallons each. Together, the total volume of fuel able to be carried and

passed to the engine through fuel lines was calculated at 2.5 gallons.

In the twin-boom design aircraft, there were instead six cylinders, with the radius,

but half the length of those used in either the pusher or conventional. They lie in three-

by-three parallel rows on each side of the engine compartment, thus resulting in

maximizing space available.

3.7 Trade Studies

3.7.1 Higher Horsepower

Trade studies were carried out regarding different engines to compare their power

outputs and fuel consumptions. A higher horsepower engine—though heavier—would

provide a greater thrust to allow the craft to take-off in a more constrained setting.

The engine used in this study was the ROTAX 912 UL 79 horsepower, four-

cylinder and four-stroke, rotary engine. Like the 582, it was also cooled via a liquid

cooling system, which contributed largely to the weight of the engine, thereby increasing

the total weight of the aircraft. According to manufacturing data regarding the 912, while

the engine performance at 3500 rpm was 35 kW, greatly outpacing the 582’s 18.5 kW at

3500 rpm, the weight of the 79 hp engine was still far too great an issue and could not be

employed using the current configurations. The power loading curve was also more linear

than the 582’s, indicating a less dependent relationship between the power and weight-to-

power ratio. [4]

19

3.7.2 Traditional Engines

A second trade study conducted was regarding the Continental O-200 four-

cylinder, air cooled, horizontally-opposed piston engine capable of reaching 100

horsepower. It, however, did not have a gear reduction system, but instead drove the

propeller directly. It weighed about 170.18 lb with no added equipment or systems, which

already exceeded the ceiling on weight distributions that the initial design had requested.

The O-200 had a fuel system that utilized 80/87 aviation gas with a fuel consumption of

6.3 at 75% operation and an average rotation rate of 2750 rpm. [5]

Though the fuel consumption for the O-200 was lower than the one given for the

ROTAX 582, the weight and lack of a reduction system posed issues that needed to be

addressed before any further progress could be made. Dealing with an air-cooled engine

with twice as much horsepower with only slightly more added weight could conceivably

be accommodated to the aerial vehicle, but not having a reduction system to limit the

speed of the propeller would be harder to handle.[5]

3.8 Future Work

Up until this point, statistical analysis had been based upon assumptions and

material provided through research. Of prime concern in the propulsion system was the

weight of the engine and the maximum outputted power. The goal, as is the case with

most aircraft, was to get as much power out of the engine without having to upgrade to a

heavier engine. Here, a point of interest lies with the liquid cooling system already

standard on the ROTAX engines. This cooling system adds to the overall weight and size

of the engine. With the size of the engine reduced, the drag created by the component

20

would decrease and allow for greater flight efficiency. Both of these issues need to be

minimized, all the while maintaining or increasing the current power of the engine. This

said, an increase in engine weight as well as horsepower would be acceptable if other

areas of the design needed to be overhauled also in order to have maximum power with

as little added weight as possible.

4. Aerodynamics (E.J.)

4.1 RFP Requirements

The initial aerodynamic design for the AUAS was primarily guided by cost

efficiency. As stated in the RFP, the system must be practical, low cost, and easily

operated. In addition, the aircraft must be able to take off from an improvised airstrip 50

ft across and must do so in less than 750 ft as well as cruise at 1,000ft AGL for transit

flights. Therefore, a configuration with a high lift coefficient and lift to drag ratio was

desired for STOL missions with desirable and efficient cruise and climb characteristics.

4.2 Airfoil Selection

From the investigation of high lift airfoils utilizing the XFOIL program and data

from the UIUC Airfoil Coordinate Database[2], the Eppler 431 airfoil was chosen as a

good initial wing choice with a sufficient balance between desirable lift characteristics

and thickness ratio for structural rigidity and internal volume. The airfoil and wing

geometry is presented in Figure 4.2.1 below.

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Figure 4.2.1 Eppler 431 airfoil cross-section

From the constraint analysis, the wing was chosen with a rectangular planform for

all three prospective configurations with geometry presented in Table 4.2.1.

Table 4.2.1 Wing geometry data for the three configurations Configuration t/c camber Chord (ft) Span (ft) Sref (ft^2) AR

Conventional Tractor 0.15 4.2% 3.93 31.3 130.8 8 Conventional Pusher 0.15 4.2% 3.93 31.3 130.8 8 Twin Boom Pusher 0.15 4.2% 3.93 31.3 138.7 8

For the initial conceptual design, the above geometry data was found to be

sufficient for performance takeoff and landing constraints, and were carried through the

entire aerodynamic analysis.

4.3 Aerodynamic Coefficients and Variables

4.3.1 CLα and CDi Calculation

Using data from the XFOIL program for the selected airfoil, a Clα curve was

generated using Reynolds and Mach numbers from performance analysis for each

mission leg at standard sea level conditions. These values are presented in Table 4.2.1

below. For concision, climb and descent mission legs are considered as part of takeoff

and landing, respectively

Table 4.3.1 Performance data for mission legs Mission Leg Velocity (fps) Mach Re Takeoff/Climb 84.4 0.075 2,099,631Cruise/Spray 81.0 0.073 2,024,645

Descent/Landing 82.7 0.074 2,067,137

22

From the Clα data, a CLα curve was produced for later use in the drag buildup and

performance analysis using the method prescribed in Raymer as Equation. 4.3.1-1

below.[1]

22

L l

l l

ARC CC C AR

α α

α α

π π

=⎛ ⎞+ +⎜ ⎟⎝ ⎠

(Equation 4.3.1-1)

Again, for brevity, the lift contribution from the horizontal tail was neglected in

this initial approximation. With values for CL, the induced drag was then calculated by

Equation. 4.3.1-2 below[1]:

2L

Dio

CCe ARπ

= (Equation 4.3.1-2)

where eo was found to be .811 by Equation. 4.3.1-3[1], which is specific to straight-wing

aircraft.

( .681.78 1 0.045 .64oe AR= − −) (Equation 4.3.1-3)

The CL values, along with the drag buildup values from the following sections were used

to produce drag polars for each configuration as shown in Figure 4.3.1 below.

23

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18

CD

CL

Tractor Pusher Tw in-Boom

Figure 4.3.1 Drag Polars for the three configurations at takeoff

4.3.2 Conventional Tractor Drag Buildup

Using the values for CL and CDi generated in the above analysis, the parasitic drag

for the first configuration (Conv. Tractor) was calculated utilizing the component buildup

method prescribed in Raymer[1] and Equation. 4.3.2 below:

( )( )

&

c

misc L P

fc c c wetDo D Dsubsonic

ref

C FF Q SC

S= +C C+∑ (Equation 4.3.2)

For this initial analysis, the drag due to losses and perturbations was neglected. For the

sake of being thorough, the value of CDL&P could be assumed to be less than 10% of the

total. The components considered for the conventional tractor configuration were the

fuselage, wing, and tail sections with miscellaneous drag components from the landing

gear, sprayer system, and engine cooling. Adding CDi and CDo together yielded the total

CD. Also, the lift to drag ratio was easily calculated as it was equal to the ratio of CL to

CD. The results of the drag buildup analysis are tabulated in Table 4.3.2 below.

24

Table 4.3.2 Conventional tractor mission segment aerodynamic coefficients (Sref=130.8ft2) Mission Segment Cl (from XFOIL) CL CDi CD L/D

Takeoff/Cruise 1.6550 1.2806 0.0805 0.1178 10.87 Cruise/Spray 1.1522 1.1006 0.0595 0.0968 11.37

Descent/Landing 1.6549 1.0760 0.0568 0.0942 11.43

4.3.3 Conventional Pusher Drag Buildup

The drag buildup for the conventional pusher configuration was calculated

following the same procedure as for the conventional tractor configuration outlined in

section 4.3.2. The components analyzed were the same as for the tractor, but with the

added drag contribution from the prop nacelle and strut. For clarity, the values of Cl

(from XFOIL), CL, and CDi calculated as in section 4.3.1 are carried over in Table 4.3.3

below.

Table 4.3.3 conventional pusher mission segment aerodynamic coefficients (Sref=130.8ft2)

Mission Segment Cl (from XFOIL) CL CDi CD L/D Takeoff/Cruise 1.6550 1.2806 0.0805 0.1183 10.82 Cruise/Spray 1.1522 1.1006 0.0595 0.0973 11.31

Descent/Landing 1.6549 1.0760 0.0568 0.0947 11.36

4.3.4 Twin Boom Pusher Drag Buildup

Again, the drag buildup for the twin boom configuration followed the same

procedure as the other two. The components considered were the two boom fuselages,

tail surfaces, wing, and engine nacelle. The miscellaneous drag components were also as

before: gear, sprayer, and engine cooling. The result of the analysis appears in Table

4.3.4 below.

Table 4.3.4 Twin boom pusher mission segment aerodynamic coefficients (Sref=138.7ft2) Mission Segment Cl (from XFOIL) CL CDi CD L/D

Takeoff/Climb 1.6550 1.2806 0.0805 0.1243 10.30 Cruise/Spray 1.1522 1.1006 0.0595 0.1033 10.66

Descent/Landing 1.6549 1.0760 0.0568 0.1007 10.69

25

4.4 Trade Studies

As part of the investigation for possible future improvements, two aerodynamic

trade studies were performed. The first was a study of the effects of high-lift devices on

CLmax and Vstall during landing. The second was analysis of the parasitic drag reduction

and subsequent performance gain from the incorporation of a Wing-Integrated Chemical

Delivery System (WICDS) over the conventional outboard sprayer. To avoid

unnecessary repetition of the same concept, the high-lift study considered only the

conventional tractor configuration.

4.4.1 Effect of High-Lift Devices on CLmax and Vstall

Considering the efficiency and durability requirements of the proposed AUAS, it

may be desirable to employ high-lift devices to increase the CLmax during takeoff, and

during landing to reduce Vstall, thereby reducing landing speed and the chance of impact

damage to the gear or other systems. Due to their structural and operational simplicity,

plain flaps and split flaps were the two possible candidates for the AUAS. Following the

method given in Raymer[1] and Equation 4.4.1-2, the change in CLmax was calculated for

both plain and split flaps. The calculation for Vstall was done using Equationn 4.4.1-2

with both the old and new values for CLmax for comparison.

max max0.9flapped

L lref

SC C

S⎛ ⎞

Δ = Δ ⎜⎜⎝ ⎠

⎟⎟ (Equation 4.4.1-1)

max

2stall

ref L

WVS Cρ∞

= (Equation 4.4.1-5)

The results are tabulated below in Tables 4.4.1-1 and 4.4.1-2. The flap effects on Vstall

are presented graphically in Figure 4.4.1.

26

Table 4.4.1-1 Conventional tractor in landing (Sref=130.8ft2 Vstall=65.39fps) with plain flaps (ΔClmax=0.9)

Sflapped/Sref del CL max

CL max (old)

CL max (flapped) CD (old)

CD (flapped)

Vstall (flapped)

(fps) % Decrease0 0.0000 1.6549 1.6549 0.09415 0.17826 65.39 0.00

0.2 0.1620 1.6549 1.8169 0.09415 0.20586 62.41 0.05 0.4 0.3240 1.6549 1.9789 0.09415 0.23605 59.80 0.09 0.6 0.4860 1.6549 2.1409 0.09415 0.26881 57.49 0.12 0.8 0.6480 1.6549 2.3029 0.09415 0.30414 55.43 0.15 1 0.8100 1.6549 2.4649 0.09415 0.34206 53.58 0.18

Table 4.4.1-2 Conventional tractor in landing (Sref=130.8ft2 Vstall=65.39fps) with split flaps (ΔClmax=1.3).

Sflapped/Sref del CL max

CL max (old)

CL max (flapped) CD (old)

CD (flapped)

Vstall (flapped)

(fps) % Decrease0 0.0000 1.6549 1.6549 0.09415 0.17826 65.39 0.00

0.2 0.2340 1.6549 1.8889 0.09415 0.21896 61.21 0.06 0.4 0.4680 1.6549 2.1229 0.09415 0.26504 57.73 0.12 0.6 0.7020 1.6549 2.3569 0.09415 0.31650 54.79 0.16 0.8 0.9360 1.6549 2.5909 0.09415 0.37333 52.26 0.20 1 1.1700 1.6549 2.8249 0.09415 0.43553 50.05 0.23

50.00

52.00

54.00

56.00

58.00

60.00

62.00

64.00

66.00

68.00

0 0.2 0.4 0.6 0.8 1

Sflapped/Sref

Stal

l Spe

ed (f

ps)

Plain Flaps Split Flaps No Flaps (clean)

Figure 4.4.1 Effect of flaps on Vstall for conventional tractor configuration From the above analysis, it is apparent that there is a more significant decrease in

Vstall for the split flaps, but also more significant increase in drag. In this application, the

plain flaps seem more desirable, since they still allow for a significantly slower stall

27

speed than the clean configuration, but they produce much less excess drag than the split

flaps. Considering also the ease of manufacture and greater simplicity (and resulting

durability) of the plain flaps, they are likely the best choice.

4.4.2 Potential Benefits of WICDS

During the drag buildup for the three configurations, it was found that the

outboard sprayer system under the wing accounted for nearly half the parasite drag. In an

attempt to gain better aerodynamic performance, a study was performed with aircraft that

incorporated a hypothetical integrated spray system that could be developed in the future.

This system, tentatively dubbed WICDS, would consist of an inboard sprayer boom and

nozzle structure that would run the length of each wing. The system would incorporate a

series of actuators that, when operated, would both slide back a narrow “door” on the

underside of the wing and turn the spray nozzles down and out to dispense the chemical.

With the WICDS, the frontal area exposed to the airflow would be drastically reduced,

which would be naturally followed by a drastic reduction in drag. Not only would the

drag contribution of the sprayer component be greatly reduced during the cruise/spray

portion of the mission, it would be virtually nonexistent during climb, turn, and descent.

The effects can be seen in Table 4.4.2 and the drag polar (for the conventional tractor in

cruise only) presented in Figure 4.4.2 below.

Table 4.4.2 CDo and L/D Data Showing the effect of the WICDS for the three configurations in cruise

Configuration CDo (conv.) CD (sprayer) CDo (WICDS)L/Dmax (conv.)

L/Dmax (WICDS)

Conventional Tractor 0.037321 0.018346 0.018975 11.67 16.37 Conventional Pusher 0.037847 0.018346 0.019501 11.59 16.13 Twin Boom Pusher 0.043823 0.016511 0.027312 10.77 13.63

28

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18

CD

CL

Conventional WICDS L/D Conventional L/D WICDS

Figure 4.4.2 Drag Polar and L/D for the conventional tractor in cruise with the WICDS. From the analysis, there is a substantial decrease in drag with the WICDS,

accompanied by a considerable increase in L/D. Utilizing WICDS would yield a

significant improvement in performance for every mission segment.

The incorporation of the WICDS would be accompanied by a loss in internal

wing volume which would limit internal configuration options, as well as increased

complexity and cost. Additionally, while the total drag contribution due to the sprayer

would be significantly lessened, there is a potential that the nozzles very near the wing’s

lower surface could perturb the airflow over the control surfaces, decreasing straight-line

stability. However, considering the dramatic change in CDo, and that the resulting

efficiency gain over the entire operating life of the vehicle would offset the increased

initial cost, the WICDS would likely be a profitable development.

29

4.5 Future Work

While the preceding aerodynamic analysis provides a good baseline estimate,

many of the calculations were based on assumptions, negations, and simplifications that

yield inaccuracy. For future work, a much more detailed lift analysis and drag buildup

will be performed for the selected configuration possibly utilizing CFD methods. Also,

the current airfoil (Eppler 431) is less than optimal, so a better airfoil will be selected

with more desirable characteristics. In addition to more thorough coefficient

calculations, the trade studies conducted show a good deal of potential for improving

performance. For the future, aggressive investigations into trailing edge devices and

integrated spray systems will be conducted.

5. Performance (R.R.)

5.1 Request for Proposal (RFP) Requirements

The RFP had many requirements that were very specific and critical to the

performance of the FUBAR-V. These RFP requirements are listed in table 5.1.1.

Table 5.1.1 RFP Requirements for the FUBAR-V

Requirements Threshold ObjectiveOperating Altitude (ft) 20 20 Takeoff Distance (ft) 750 600 Landing Distance (ft) 750 600

Other RFP requirements include the operating speed being 1.3 times the stall

speed of the aircraft, carrying out short ferry flights of 2 miles at a 1000 foot altitude and

also carrying 20 minutes of reserve fuel. The three chosen aircraft designs were mainly

based on the above given RFP requirements.

30

As an unmanned agricultural aircraft, one of the major performance areas focused

upon was the actual way to spray the chemicals over the entire field. Different flight

paths were analyzed and the most efficient way to spray the chemicals was found. Other

usual performance calculations, based on the RFP requirements, including turn radius,

turn rate, velocity which requires minimum power, velocity for the best rate of climb,

stall velocity of the aircraft, maximum range, wing loading, and power loading were

calculated as well.

5.2 Mission Profiles

The first mission that the aircraft had to perform was 1,2) to takeoff from a 50

foot wide grass airstrip, 3) climb to an altitude of 50 feet, 4) descend to 20 feet to the

field, 5) spray chemicals over the entire field, 6) ascend to 50 feet, and 7,8) finally land.

This is outlined in the Fig 5.2.1.

Figure 5.2.1 First Mission Profile

The second mission was 1) to takeoff, 2) climb to 1000 feet, 3) cruise for 2 miles,

4) descend and 5) finally land as shown in Fig 5.2.2.

3

2 4

1 5

31Fig 5.2.2 Second Mission Profile

5.3 Aircraft Parameters

The performance of the FUBAR-V relied mainly on the configuration, propulsion

and aerodynamic characteristics. The aircraft parameters from these areas, which were

used in the performance calculations of the three aircraft designs, are shown in Table

5.3.1.

Table 5.3.1 Aircraft Parameters Used

Conventional

Tractor Twin Boom

Conventional Pusher

A/C Gross Weight(lbs) 1100 1166 1100 Total wing area(ft2) 130.8 138.7 130.8

Max fuselage width(ft) 2 2.6 2 Max fuselage height(ft) 2 4.33 2 Propeller diameter(in) 62 62 62 Propeller efficiency 0.9177 0.9177 0.9177

Cl max 1.55 1.55 1.55 Max horsepower(bhp) 65 65 65

Max prop RPM 3525.48 3525.48 3525.48 Engine bsfc 0.41 0.41 0.41

Max fuel (gal) 2.5 2.5 2.5 Reserve fuel(min) 20 20 20

Drag coefficient(Cdo) 0.037 0.044 0.038 Wing loading 9 9.54 9

As the three aircraft designs were generally similar to each other, the parameters

used were also the same, except for a few factors like the drag coefficient and the weight.

The change of drag coefficient parameter in the twin boom and the conventional

aircraft designs would affect the cruise velocity and the fuel fraction, since the fuel

fraction increases and the cruise velocity decreases as the drag coefficient decreases.

From the aircraft parameters above, the performance constraints shown below in

Table.5.3.2 were estimated mainly using the formulas from Raymer.

32

Table 5.3.2 Estimated Performance Parameters

Conventional

Tractor Twin boom

Conventional Pusher

V max (knots) 50 50 50 V best ROC (knots) 48 49 48 Vmin power (knots) 47 48 47 V stall, clean(knots) 41 43 41 Max range (miles) 56 52 56

Power loading (lbs/hp) 16.92 17.94 16.92

5.4 First Mission

5.4.1 Takeoff and Initial Climb

According to the RFP requirement, the aircraft should takeoff within a 750 foot

takeoff distance in a grass improvised airstrip, which is 50 feet wide. Also since the

altitude to which the aircraft had to climb was 50 feet, the equations and calculations [6]

for the 50 foot takeoff height clearance were used.

Since the airstrip was improvised grass, the coefficient of rolling friction was

chosen as 0.05. The takeoff velocity was determined as 50 knots, which is the estimated

maximum velocity the aircraft could fly, while taking into consideration the engine

capacity and payload. The thrust at 0.7 takeoff velocity was calculated, and then using

this value the takeoff ground roll distance was calculated. The takeoff distance for the

conventional type aircrafts was 386 ft, and for the twin boom aircraft was 409 ft.

After takeoff, the distance required for the aircraft to climb to 50 feet altitude was

247 feet for the conventional aircraft designs and 260 feet for the twin boom aircraft, at

49 knots which was the velocity calculated for the best rate of climb. The time to climb

was calculated as 3.14s for the twin boom aircraft design and 2.98s for the conventional

aircraft designs. The profile of the takeoff and climb is shown in Fig 5.4.1.

33

Fig 5.4.1 Takeoff and Climb Profile

The total fuel consumed for the aircraft to takeoff and climb to 50 feet is 0.42

liters. In these above mentioned cases, the twin boom aircraft weighed slightly heavier

than the other two conventional aircraft designs, which increased the wing loading and in

turn takeoff and the climb distance increased. This clearly indicated that the takeoff and

the climb performance worsened as the wing loading increased.

5.4.2 Descent

The descent segment of the mission did not contribute significantly towards

defining the critical characteristics and the parameters that would affect the performance

of the aircraft. Due to this reason, all the calculations on this decent segment were based

on historic values obtained from Raymer [1].

The fuel weight fraction was taken as 0.995 as the descent was from a 50 foot

altitude to the cruise altitude of 20 feet. The fuel consumed from this mission segment

was 0.045 liters.

5.4.3 Fertilizers Spraying Segment

As mentioned before, the purpose of this mission was to fly the aircraft over the

agricultural field and spray the fertilizer chemicals efficiently from an altitude of 20 feet.

34

The main challenge was to find an efficient way to fly around the field. Many different

flight paths were considered and the most efficient path was chosen from them. Three of

these flight paths are shown in Fig 5.4.3:

Fig 5.4.3 a) S shaped turns b) Spiraling into centre c) Translating rectangular

ellipse

According to the RFP, the field is 0.5 miles to 1000 feet, and has to be sprayed

with fertilizers, which is considered as mission’s cruise segment including appropriate

turns. The first flight path shown in Fig 5.4.3 a) was efficient as the whole field was

sprayed evenly with the spray width of 32.55 feet. But unfortunately the aircraft had to

make 22, 180° turns, and this proved to be a very difficult maneuver for the aircraft to

perform. To prevent the complication of the 180° turns, the idea of a spiral flight path

moving towards the centre was formed. However, when the aircraft got closer to the

centre of the field, the turns became very sharp, which again proved to be too complex a

maneuver.

To overcome all these problems, the translating rectangular elliptical flight path

was chosen, allowing the aircraft to easily perform blunt sustained turns. The turn

calculations for this flight path for all the three aircraft designs are summarized in Table

5.4.3.

35

Table 5.4.3 Fertilizer Spray Calculation Summary

Wingspan (feet) 31 Spray width(feet) 32.55

Length(feet) 2640 Width(feet) 1000

Number of passes 16 Total Distance(miles) 18.93Chosen Load factor 5 Turn Radius(feet) 39.78

Minimum turn radius (R min)(feet) 35.53

The calculation of fuel consumed for this mission segment was slightly different

to the other mission segments, as fertilizer spraying was considered as a single payload

drop, rather than a gradual reduction of payload from the tank. This helped to alleviate

the complexity with which these performance calculations were subject to. This reduced

the wing loading of the conventional aircraft designs to 6.54 lbs/ft2, from the initial

estimate of 9 lbs/ft2, as the weight of the aircraft decreased by 300lbs after the spraying

process. Using this wing loading, the lift-to-drag ratio was calculated and the fuel weight

fraction for this mission segment was found as 0.562 from the Breguet range equation [7].

From this fraction, the fuel consumed for this segment was calculated as 3.94 liters.

For the twin boom aircraft design, the wing loading was reduced from 9.54 lbs/ft2

to 7.09lbs/ft2. Using the same procedure as above the fuel weight fraction was found as

0.595 and the fuel consumed was 3.63 liters. This proved that the twin boom aircraft

consumed less fuel for this mission segment compared to the conventional aircraft

designs. This is mainly due to the slightly bigger drag coefficient in the twin boom

aircraft. This also confirmed the fact that, as the drag coefficient increased the fuel

fraction increased consequently.

36

5.4.4 Second Climb

In this climb segment the aircraft had to climb from 20 feet to 50 feet. As used in

the first climb, the aircraft velocity for the best rate of climb was 48 knots was used. The

fuel fraction for this climb segment was calculated as 0.989. From this fraction the fuel

consumed was 0.055 liters. The total consumption for this mission segment remains the

same for all the three aircraft designs.

5.4.5 Descent and Landing

Just like the takeoff and the initial climb calculations, the equations for the 50 foot

landing height clearance were used to calculate all the parameters for this mission

segment. The approach velocity usually is 1.3 times the stall speed [8], was coincidently

the same as the RFP constraint, so the velocity was kept constant at 48 knots. The fuel

fraction for the descent from 50 feet was estimated historically as 0.995 from Raymer [1].

The fuel consumed for just this descent was 0.025 liters.

The touch down velocity was calculated as 48.3 knots, which is 1.15 times the

stall velocity of 42 knots [8]. Using this velocity the landing distance was calculated as

462 feet for the conventional aircraft designs and 489 feet for the twin boom

configuration. The slight difference between the landing distances for each aircraft design

was because the twin boom aircraft had a stall speed of 42 knots compared to the 40

knots of the conventional aircraft designs. The landing distance was well within the RFP

constraint of 750 feet. The fuel fraction for landing and taxi-back segment was calculated

as 0.993 and in turn, the fuel consumed was found as 0.035 liters.

37

5.5 Short Ferry Mission

The takeoff calculations were the same as the first mission, so the takeoff distance

was 386 feet and 409 feet for the conventional aircraft and the twin boom aircraft design

respectively. The fuel consumed was also the same as before, 0.42 liters. Climb to takeoff

to 1000 feet, was again done at the velocity for the best rate of climb, 49 knots. The fuel

fraction for this segment was calculated as 0.866 and the fuel consumed was 1.64 liters.

The cruise for 2 miles at 1000 feet was done at the velocity for minimum power at 48

knots and the estimated fuel fraction for this segment was 0.948 as well as the fuel

consumed was 0.40 liters. Since the descent did not have any significant contribution

towards the performance of the aircraft, the distance covered was added to the cruise

range. Finally, the landing distance was also similar to the first mission, 439 feet for the

conventional aircraft and 466 feet for the twin boom aircraft. The fuel consumed also

remained same as the first mission, at 0.035 liters.

5.6 Summary and Looking Ahead

All the three aircraft designs met the performance RFP constraints, including the

takeoff and landing distances, stall speed constraint and the short ferry flight constraint.

Also the three aircraft designs were capable of flying longer, given the fact that the total

amount of fuel consumed by the conventional aircraft designs was 4.48 liters out of the

9.54 liters available. Even the twin boom aircraft consumed only 4.18 liters out of the

9.54 liters available. This remaining fuel also included the reserve fuel for 20 minutes.

Overall, the differences in performance characteristics of the three chosen aircraft designs

were very little, while the twin boom aircraft consumed lesser fuel than the other

conventional designs. This needs to be confirmed with more accurate calculations, as

38

fairly crude calculations were carried out for the mission analysis. Especially, the climbs

and the descent calculations have to be done more exactly, as they are mostly calculated

here with historic data.

One other area to concentrate in future is the fertilizer spray mission segment,

where the gradual payload reduction system must be introduced, and the performance

calculations must be carried out accordingly. Further work should also include the section

about the aircraft performance at higher altitudes and with heavier payload.

6. Stability and Control (G. F.) 6.1 Introduction

Stability and control of an aircraft is vital to the performance of the aircraft and

can be considerably linked to its aerodynamic characteristics. The RFP set forth a

guideline for a low cost and easily operational agricultural unmanned aerial vehicle

(UAV). Controllability of aircraft is vital to any design and this RFP calls for no

difference. Stability analysis was done to determine the sizes of the vertical and

horizontal stabilizers as well as the rudder and elevator control sizes for their respective

stabilizer. To comply with the RFP, this aircraft will require a Short Take Off and

Landing (STOL), consisting of 750 ft in length and 50 ft wide improvised grass or gravel

runway. Design must also be able to transport and disperse 235 lbs liquid or 300 lbs solid

agricultural payload to be distributed throughout a 2640 ft by 1000 ft rectangular field.

This dispersion of the payload will directly affect the stability and control of the aircraft

by changing the value of the Center of Gravity (CG) of the aircraft. Other RFP

requirements will be described in their appropriate sections.

39

6.2 Initial Stabilizer Sizing

Initially many tail designs were suggested for the three proposed designs. The

conventional tractor design that was proposed will consist of a conventional tail design.

Another concept of taking the conventional tractor and turning it into a pusher model was

considered and will be given the same conventional tail design given to the tractor. A

twin boom design will complete the three proposals and this will call for an

unconventional design that can be viewed in the three view provided in this report.

Each calculation will be very similar throughout each of the previously mentioned

designs. For a first order approximation the center of gravity of the wings will be

considered to be located at the wings aerodynamic center, approximately the quarter cord

of the wing. The quarter cord of the wing will be set at a location 3ft aft of the nose in

the twin boom design and 6.75ft, 6.5ft aft of the nose for the tractor and pusher

respectively.

40

Table 6.2.1 Tail Size

Conventional Tractor Conventional Pusher Twin Boom

ARvt 4.5 4.5 4.5

ARht 2.75 2.75 2.75

Svt 13.33953 ft2 13.33953 ft2 13.90737 ft2

Sht 20.52235 ft2

20.52235 ft2

21.59714 ft2

bvt 4.864864 ft 4.864864 ft

4.96733 ft

bht 9.60992 ft 9.60992 ft 9.858354 ft

lt 11.7697 ft 11.7697 ft

12.07397 ft

C 3.923234 ft 3.923234 ft 4.024656 ft

tC 2.135538 ft 2.135538 ft 2.190745 ft

actX 8.77999 8.77999

8.558731

To find the values of Svt and Sht a relationship between the aerodynamic

characteristics of the aircraft were used, these Equations, 6.2.1 and 6.2.2, where used to

calculate the tail area. [1]

vt

vtvt l

bScS = Equation 6.2.1

41

ht

htht l

SCcS = Equation 6.2.2

Where cvt and cht are equal to 0.03 and 0.50 respectively, these values were taken from

historical data from Table 6.4 Raymer.[1] Also since lvt and lht are unknown, in equations

6.1 and 6.2, historical data where used as for these calculations where lht and lvt are 60%

of fuselage length for the tractor configuration and 45% of fuselage length for the pusher

models. [1]

6.3 Tail Configuration

A variety of tail designs were obtainable for each of the three designs presented.

Two of the vehicles, the conventional tractor and the conventional pusher, will be flying

the same conventional tail design. The third, twin boom vehicle, will consist of two

vertical tail segments, one mounted on each boom, and one horizontal tail component

mounted to the top of each vertical segment. Both tail designs can be viewed in Figure

6.3.1 and 6.3.2 respectively. Also for more tail sizing information refer back to Table

6.2.1 above.

Figure 6.3.1Figure 6.3.2

6.3.1 Twin Boom Tail

Calculation for the vertical tail span, bvt, was calculated assuming one continuous

vertical tail as found in standard conventional then split into two halves for the two

42

vertical segments of our twin boom tail design. This concept gave us the needed vertical

tail dimensions needed for stability and control, but also helped keep the vertical size of

the tail lower.

6.3.2 Conventional Tail

The conventional pusher and tractor design called for a tail approximately the

same specification of the twin boom design described in section 6.3.1, but we opted for

an easy to construct and maintain conventional tail design as well. Crop dusters do not

require a lot of maneuverability like your modern military jet fighter would need in

battle, for this fact this conventional “t-tail” design will give enough stability to allow for

the operations at hand.

6.4 Control Surfaces

Sizing for control surfaces came strictly from historical data at this point in

design. Values of ailerons and rudder sizes were obtained from historical data located in

Raymer. [1] Located in Table 6.5 in Raymer elevator size, Ce/C, and rudder size, Cr/C,

were taken as values from sailplanes. [1] Our designs will be taken as 0.43 for elevator

sizing and 0.40 for rudder size. Located in Table 6.4.1 are values comparing the two

control surface sizing and its associated position within the plane.

43

Table 6.4.1 Control Surface Sizing

Conventional Twin Boom Location

Span

(ft)

Chord

(ft)

Span

(ft)

Chord (ft)

Rudder Vertical Tail 3.895913

0.800827

4.080809

0.82153

Elevator Horizontal Tail 4.324464

4.05

8.872519

4.05

Ailerons Wing 6.199878

0.395074

6.524578

0.405288

6.4.1 Rudder Configuration

Rudder configuration for the conventional tail will be placed at the trailing edge

of the vertical tail fin. Historical data gives the value for rudders to be between 25-50%

of the tail chord. Given this large range and this being an initial design stage, a value of

37.5% was chosen as the average between the two extremes given above. The rudders

present on the conventional tail design will consist of no taper because the tail is an un-

tapered airfoil, at least assumed to be in this initial design.

On the twin boom design the rudder will be given the same 37.5% value chosen

from the average of the historical data taken from Raymer, but will be split into two equal

segments distributed between the two vertical parts of the tail. [1] The same consideration

on taper will take place for the twin boom design.

44

6.4.2 Elevator Configuration

Elevators will be placed on the vertical tail surface of each tail configuration. For

the conventional tail the elevators will be have its span divided in half for the

conventional tail design, with each horizontal tail segment taking on half of the elevator

role. For the twin boom this will not be necessary because of the large horizontal,

continuous airfoil section that will have the elevators installed on the trailing edge.

6.5 Trade Studies

6.5.1 Trade Study One

A major design input of any aircraft will be the center of gravity (cg) of the

aircraft. For the models above the static margins of all the aircraft were relatively large

compared to historical data, the first study will be to reconfigure the locations of the

internal components. First by moving the payload, and its corresponding units into a

different configuration, lead to the center of gravity, on the twin boom design, to move

forward, away from the nose, from 9.8393 to 9.3625.

Table 6.5.1 Trade Studies

Original CG Trade Study CG

Location 9.8393 ft 9.3625 ft

SM difference 0.382587 0.398978

6.5.2 Trade Study Two

On the conventional design a trade study was done to vary the tail size, which in

turn varied the length of the tail needed to compensate for the moments. Increasing of

the tail size required a smaller tail length to be used to correct the moment for stable

45

flight. The opposite is true also, decreasing the size of the tail sections will require a

longer tail length, lt, to compensate for the decreased size in tail.

6.6 Future Work

For the future of this design will call for changes to be made in its center of

gravity and/or neutral point. This will accompany changes in the static margin and make

the plane either more stable or less stable depending on the desired mission.

Maneuverability has a direct correlation to the stability of the aircraft. If a more

maneuverable crop duster is needed for either more non-conventional dusting, then a

smaller static margin will be needed. For example the F-22 Raptor has a static margin of

0 to -15% for its maneuverability, and requires a “relaxed static stability (RSS)”

computerized control system. [1]

Also a military version could be created with a change in power plant and static

margin. This could lead to a wider application range and production. If a military role

was to be adapted for reconnaissance a V tail could be placed on the conventional design.

For this design a reduction in weight would allow for a cheaper operation and

production costs. This would also call for a major redesign on the materials used on the

application. For instance replacing the aluminum skin with something of lighter weight

that does not jeopardize such as a composite or lighter alloy will allow for a smaller

engine to run the aircraft or a higher thrust to weight ratio. A higher thrust to weight ratio

would benefit both the military application and allow for more fertilizer to be carried for

the agricultural application. Most importantly however is to select the proper design for

our application and continue that design into the following semester.

46

7. Structures (P.M.)

The requirements of the RFP indicate that the crop duster will be launched from

make-shift runway and will have to endure rugged conditions. It also requires that the

design be simplistic, reliable, and affordable to all nations. These criteria will be kept in

mind throughout the structural analysis of the FUBAR-V designs. This analysis includes

formation of a V-n diagram, load path considerations, material selection, and selection of

the landing gear.

7.1 V-n Diagram

The V-n diagram is essential for determining the proper structural design and

components of the aircraft necessary to maintain loads encountered in flight. These loads

can be produced by many sources. Two prominent ones are produced by lift during high-

g maneuvers and wind gusts and are known as maneuver loads and gust loads

respectively. A V-n diagram based on the Eppler 431 airfoil is found in Figure 7.1.1.

47

-3-2-1012345

0 10 20 30 40 50 60 70 80 90 100

V (mph)

n

Positve Gust Load at Vc Negative Gust Load at Vc

Positive Gust Load at Vd Negative Gust Load at Vd

High AOA Max q

Gust Load btw Vc and Vd Gust Load btw Vc and Vd

Maneuver Envelope

Figure 7.1.1 V-n Diagram for Wing Loading of Nine

Some important features of the diagram are highlighted in the legend. The “High

AOA” point represents the lowest velocity (without stall) at which the airplane can

encounter the maximum load factor. The “Max q” point indicates the speed at which the

maximum dynamic pressure is encountered by the aircraft. This speed is referred to as

the dive or maximum speed of the crop duster. [1] A cruise speed of 65 mph was used for

the calculations. This value is approximately 10 mph higher than the value at the design

point (using the value at the design point produces a diagram in which the velocity at

point AOA is higher than the dive speed which is incorrect). This is a discrepancy that

needs to be resolved in future work. A possible solution to this problem is finding

methods to lighten the aircraft in order to increase the cruise velocity. However, since

the values are reasonably close this V-n diagram will be a good approximation of the

final one.

48

Regulations from the FAR Part 23 were used in determining the limit load factors

and dive speed on the V-n diagram. In accordance with FAR Part 23 Sections 335 and

337[9], the maximum positive load factor is given by Equation 7.1.1.

n+=2.1+(24,000/(W+10,000)) (Equation 7.1.1)

The maximum negative load factor is given by Equation 7.1.2.

n-=0.4*n+ (Equation 7.1.2)

Finally Equation 7.1.3 gives the dive speed based on the cruise speed.

Vd=1.5*Vc (Equation 7.1.3)

After calculating these parameters it is possible to determine the positive and negative

gust loads using Equation 7.1.4.

ng±=1±(K*U*a*V)/(498*(W/S)) (Equation 7.1.4)

According to FAR Part 23, U equals 30 ft/s for cruise velocity and 15 ft/s for dive

velocity. K is the gust alleviation factor which takes into account the fact that heavier,

larger planes encounter gusts at slower rates than lighter, smaller aircraft. The slope of

the normal coefficient curve is given by a. Values used in the V-n diagram creation are

tabulated in Table 7.1.1. These values are the building blocks in determining material

selection, fuselage construction, and proper part mounting.

Table 7.1.1 V-n Diagram Data W

(lb) (W/S) Vc

(mph) Vd

(mph) a ng+ at

Vcng- at

Vcng+ at

Vdng- at

Vdn+ n-

1,100 9 65 97 6.2979 2.77 -0.77 2.32 -0.32 4.26 -1.70

Referring to Figure 7.1.1, there are two gust loads which my produce a load factor

outside of the maneuver envelope. One is a positive load factor (at Vc) and the other is a

negative load factor (at Vd). Since the structural components of the aircraft will be built

49

to withstand loads contained along the maneuver envelope, it is vital that all gust loads be

contained within this envelope. Thus the envelope must be enlarged to include these

loads. This may result in the aircraft requiring a more expensive, stronger material which

in turn may lead to an increase in weight and cost. However, this is a better alternative

than structural failure during flight.

In order to better understand how wing loading affects the V-n diagram, a trade

study was conducted. The weight was held constant at 1,100 lbs while the wing area was

increased so that (W/S) decreased to a value of 4. The V-n diagram for this variation is

given in Figure 7.1.2.

-4

-2

0

2

4

6

0 10 20 30 40 50 60 70 80 90 100

V (mph)

n

Positive Gust Load at Vc Negative Gust Load at Vc

Positve Gust Load at Vd Negative Gust Load at Vd

High AOA Max q

Gust Load btw Vc and Vd Gust Load btw Vc and Vd

Maneuver Envelope

Figure 7.1.2 V-n Diagram for Wing Loading of Four

50

The maneuver envelope is considerably larger for the smaller wing loading. The

“High AOA” point is reached at around 30 mph less than for plane with a wing loading

of 9. This shows that planes with more reference area experience greater loads at slower

speeds. Another notable feature of Figure 7.1.2 is that there still exist two gust loads that

may cause the plane to experience load factors outside of the envelope. In fact, the whole

range of negative gust loads encountered from Vc to Vd (the green line) may produce

load factors outside of the operating envelope. In contrast to this, the range of positive

gust loads is completely contained with the envelope. An important conclusion is that as

the wing loading decreases special attention needs to be placed on the negative gust

loads. Based on the previous two diagrams, if the wing loading was raised to a higher

value (say 15) it is likely that the entire range of positive gust loads would be outside the

envelope (while the negative ones would be contained within the envelope).

Construction of this diagram will be saved for future work.

7.2 Load Paths and Structural Load Calculations

The loads to be encountered by the aircraft have been determined from the V-n

diagram. It is now necessary to determine how these loads will be carried through the

structure. Shear and bending moment diagrams at the wing root also need to be

constructed.

The modern conventional construction of the fuselage is known as the monocoque

design. This design does not rely solely on internal members to support loads, but rather

on the strength of the skin covering the wings and fuselage. There are three variations of

this fuselage design: true monocoque, semimonocoque, and reinforced shell. In the true

51

monocoque configuration, no bracing members are present throughout the skin and it will

hold a majority of the load encountered by the airplane. This can lead to very high

weights and cost since ample material must be used to ensure the strength of the skin.

Since low cost is of upmost concern in the RFP, the true monocoque variation will not be

used. In both the semimonocoque and reinforced shell variations, the airplane skin is

reinforced which allows for less skin material to be used. The reinforced shell uses a

complete framework of both lateral and longitudinal structural components, while the

semimonocoque uses only longitudinal members. Once again, to keep costs and weight

at a minimal the semimonocoque variation will be used. An example of a

semimonocoque fuselage configuration is given in Figure 7.2.1.

Figure 7.2.1 Semimonocoque Structure Variation[10]

The airplane in Figure 7.2.1 has a similar layout to the conventional tractor design and

can be studied to get a general idea of how the internal structure of the FUBARV designs

will be modeled. The internal structure is primarily composed of stringers, longerons,

52

and bulkheads. The longerons hold the bulkheads, which are used at high concentration

load points such as the wing-fuselage juncture, in place. The bulkheads support the

stringers, which provide shape and stability to the airplane skin. A more detailed analysis

of the internal configurations, including sketches and modeling, will be saved for future

work.

The shear and bending loads encountered by the wing root constitute a high

concentration load that the bulkhead must maintain. Thus it is of particular interest to

find these load values. Schrenk’s approximation was used to obtain the lift per unit span

for the rectangular planform used on the FUBARV designs. This data was plotted in

Excel and an equation for the lift per unit span (in terms of distance from the wing root)

was found. This equation was integrated and multiplied by the dynamic pressure in order

to obtain the total wing lift for the critical cases of high AOA and max q. The weight of

the wing was estimated using Table 15.2 in Raymer which states that each square foot of

wing weighs approximately 2.5 lb. With the wing lift, wing weight, and load factor

(taken from V-n diagram) now known, simple beam analysis can be performed to

calculate the shear force and bending moment at the wing root. Results for the two

critical cases are found in Table 7.2.1.

Table 7.2.1 Shear and Bending Moment Loads at Wing Root for Limiting Cases

Critical Case Lift (lb) n+Wing weight

(lb) Shear force at root (lb)

Bending moment at root (lb*ft)

High AOA 1,319 4.272 152.3 4,984 18,083 Max q 2,942 4.272 152.3 11,918 43,240

These results will be essential in determining what kind of material to use. If the

structure fails at the wing-fuselage juncture, then obviously the whole airplane fails.

Thus it is vital that a strong, durable material capable of tolerating these loads is used.

53

7.3 Material Selection

The material needed for this RFP must be resistant to corrosion and the elements,

strong, of low cost, and readily available. There are four widely used materials that were

considered for the structural components on the aircraft. These four materials are wood,

aluminum, steel, and titanium.

Two of these materials were eliminated immediately on the grounds that they did

not match the requirements in the RFP. The first of these was wood. Though cheap and

available anywhere in the world, wood would rot too easily in the harsh conditions

experienced by a crop duster. Titanium has the highest strength-to-weight ratio of the

four materials mentioned and is also durable. However, it is also the most costly and

difficult to produce so it too was ruled out.

The two remaining materials best fit the needs of the RFP. Both are in wide use

and have enough strength to withstand loads encountered by the crop duster. Steel is

around one-sixth the cost of aluminum. However, aluminum is especially resistant to

corrosion. It is because of its excellent resistant to the elements that aluminum was

chosen as the main material used in construction of the aircraft. Though the initial cost

may be slightly higher using aluminum, the material’s resistance to chemical degradation

should save money in the long run with less repairs and maintenance needing to be

performed on the aircraft. Aluminum is often strengthened through alloying with other

metals (aluminum 2024, aluminum 7075). However, this also makes the metal more

susceptible to corrosion. Since aluminum 2024 is used on the crop duster, it would be

wise to coat it with pure aluminum to regain resistance to the elements. Future work in

54

this area will include detailed calculations to ensure the structural stability of the

fuselage-wing juncture.

7.4 Landing Gear

Proper landing gear design is essential to the structural integrity of the aircraft.

Each landing gear configuration comes with its own set of advantages and disadvantages

which must be weighed accordingly before a final configuration can be chosen. After the

configuration is decided upon, the location of the landing gear on the airplane (relative to

the center of gravity) can be determined by following specific guidelines that ensure the

stability of the aircraft. After the location of the landing gear is known, the type of tire(s)

used can be determined. The tires hold the entire weight of the aircraft during takeoff,

and during landing they may hold additional dynamic forces in addition to the static

weight of the aircraft. The tires must be able to maintain these loads and also deal with

the extra stress created by the rough, makeshift runways that the aircraft will be utilizing.

The three configurations for the FUBAR-V crop duster each have a different

landing gear configuration. The conventional tractor uses a taildragger configuration,

while the conventional pusher uses a single main configuration, and the twin-boom

pusher design uses a slightly modified single main configuration. These configurations

may be modified or changed in the future to accommodate updated design parameters.

7.4.1 Configuration for Conventional Tractor

The conventional tractor crop duster is best suited for a taildragger configuration.

This configuration gives the plane the proper propeller clearance necessary (at least nine

inches).[1] An additional advantage is that since the nose of the airplane is tilted up, the

airplane’s angle of attack at takeoff is increased and this allows the wings to generate

55

more lift (thus reducing takeoff distances). This is especially helpful since the plane will

most likely always be taking off from makeshift runways of no specified length. This

landing gear would have also been used on the other configurations if not for its inherent

unstable nature. Since the center of gravity lies behind the two main gears, the pilot must

ensure that the two main wheels touch the ground at the same time. If one wheel touches

the ground before the other, the location of the center of gravity may cause the plane to

veer in the direction of this wheel. This can cause the plane to enter into a highly banked

ground loop, possibly resulting in a wing tip scraping the ground. [11] Since the spray

system is attached to the bottom of the wing this is a highly undesirable situation in

which the system may be damaged or rendered inoperable. There are certain

specifications that can be followed, however, in order to reduce the stability issues with

the taildragger configuration.

There are three main guidelines that should be followed in order to maximize

stability. The first is that the tail down angle (i.e. the amount the angle of attack is

increased at take off) should be in-between 10 and 15 degrees. Also, the center of gravity

should be located at an angle of 16-25 degrees from the vertical of the main wheels.

Finally, as viewed from the rear take off position, the main wheels should be laterally

separated at an angle of greater than 25 degrees from the vertical line passing through the

center of gravity. Graphical descriptions of these requirements can be found below.

56

Figure 7.4.1 Stability Requirements for Taildragger Configuration[1]

If a taildragger configuration was used on the other designs of the crop duster and

these regulations were followed, the chances of the aircraft encountering the ground loop

would be greatly reduced. However, the spray system is the most vital part of the aircraft

(without it the plane is no longer a crop duster) and its integrity should not be risked

unnecessarily. The taildragger was necessary in order to provide proper clearance for the

props on the tractor design. However, the engine on the pusher configuration will be

mounted on top of the central nacelle and thus propeller clearance will not be an issue. A

more stable configuration can be used in order to protect the spray system.

7.4.2 Configuration for Conventional Pusher and Twin boom Pusher

This more stable configuration comes in the form of the single main gear for the

conventional pusher design. This design eliminates the possibility of ground looping and

is both simple and lightweight. Historically, it has been used on lighter aircraft (a crop

57

duster being a good example of a light aircraft). In addition, this configuration also

produces less drag than the taildragger. The twin boom pusher design follows the single

main gear configuration with a slight modification. The central nacelle does not extend

as far aft as the twin booms because the engine is mounted on the backside of the nacelle.

Since the tail consists of two vertical stabilizers (one from each boom) attached on top by

a horizontal stabilizer, there is no central location to which the tail wheel could be

attached. Thus each boom will have a tail wheel extending from its underside. This will

increase the overall weight and drag of the airplane slightly, but the effects should be

minimal on the performance.

7.4.3 Tire Selection

After determining the gear configuration used, the tires on the gears must be

properly sized in order to support the weight of the aircraft. The width and diameter of

the tires was determined by using data in Table 11.1 in Raymer. The values found using

this data were multiplied by 1.30 in order to account for the aircraft operating on rough,

unpaved runways. For the taildragger configuration, it was assumed that the two main

tires combined carry 90 percent of the total weight of the airplane while the tail tire

carries the remaining 10 percent. For the single main gear configuration, it was assumed

that the main central tire carries 70% of the load while the other wheels carry 10 percent

each. For the modified configuration on the twin boom pusher, the two tail tires are

assumed to each carry 5 percent of the total load (equal to the 10 percent carried by the

single tail tire on the conventional pusher). The diameter and width of the tires for each

configuration are tabulated in Table 7.4.3.

58

Table 7.4.3 Tire Width and Diameter

Landing Gear Configuration

Main Tire

Diameter (in.)

Main Tire

Width (in.)

Wing Tire

Diameter (in.)

Wing Tire

Width (in.)

Tail Tire

Diameter (in.)

Tail Tire

Width (in.)

Taildragger 16.489 6.2308 N/A N/A 10.124 4.029 Single Main

Gear 19.9667 7.3932 10.124 4.029 10.124 4.029

Modified Single Main

Gear 19.9667 7.3932 10.124 4.029 7.949 3.245

From Table 11.2 in Raymer, a tire size of 7.00-8 will be used for the main wheels,

and a tire size of 5.00-4 will be used for the auxiliary (wing and tail) wheels.[1]

Researching the availability and cost of these tires from different manufacturers will be

considered in the future. The specific shock absorber to use on the landing gear will also

be determined at a later date. This will allow a