CONCEPTUAL DESIGN REPORT Agricultural Unmanned Aircraft System (AUAS) Team Two-CAN Team Member Area of Responsibility Albert Lee (Team Leader) Aerodynamics Chris Cirone Performance Kevin Huckshold Configuration Adam Kuester Structures Jake Niehus Stability and Control Michael Scott Propulsion AE440 Senior Design November 16, 2007
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CONCEPTUAL DESIGN REPORT
Agricultural Unmanned Aircraft System (AUAS)
Team Two-CAN
Team Member Area of Responsibility
Albert Lee (Team Leader) Aerodynamics
Chris Cirone Performance
Kevin Huckshold Configuration
Adam Kuester Structures
Jake Niehus Stability and Control
Michael Scott Propulsion
AE440 Senior Design
November 16, 2007
i
TABLE OF CONTENTS
Nomenclature iii
Executive Summary (AL) v
1. Introduction (JN) 1
2. Configuration Selection (KH) 3
2.1. Objectives 3
2.2. Morphology 3
2.3. Configuration Selection 4
2.4. External Configuration 6
3. Sizing Analysis (AK, CC) 10
3.1. Initial Sizing (AK) 10
3.2. Constraint Analysis (CC) 11
4. Performance (CC) 13
4.1. Introduction to Performance 13
4.2. Takeoff Analysis 13
4.3. Cruise Analysis 15
4.4. Mission Time 19
4.5. Fuel Estimation 20
4.6. Trade Studies 20
4.7. Future Work in Performance 21
5. Aerodynamics (AL) 22
5.1. Introduction to Aerodynamics 22
5.2. Airfoil Selection and Wing Geometry 22
5.3. Lift, Drag, and Efficiency Analysis Method 26
5.4. Lift, Drag, and Efficiency Results 27
5.5. Future Work in Aerodynamics 29
6. Propulsion (MS) 30
6.1. Introduction to Propulsion 30
6.2. Methodology 30
6.2.1. Engine Selection 30
6.2.2. Propulsion System Selection 30
6.2.3. Engine Air Intake 32
6.2.4. Fuel System 33
6.3. Results 33
6.3.1. Engine Selected 33
6.3.2. Propulsion System Selected 34
6.3.3. Engine Air Intake 36
6.3.4. Fuel System 36
6.4. Future Work in Propulsion 37
7. Stability and Control (JN) 38
7.1. Introduction to Stability and Control 38
7.2. Tail and Control Surface Sizing 38
7.3. Longitudinal Static Stability and the Static Margin 41
7.4. Future Work in Stability and Control 44
ii
8. Structures (AK) 45
8.1. V-n Diagram 45
8.2. Materials Selection 47
8.3. Fuselage Structure 47
8.4. Wing-Fuselage Attachments 48
8.4.1. Structure at Wing-Fuselage Interaction 48
8.4.2. Wing Root Limiting Loads 49
8.5. Landing Gear 50
8.5.1. Landing Gear Structure 50
8.5.2. Landing Gear Location 51
8.6. Future Work in Structures 52
9. Configuration (KH) 53
9.1. Introduction to Configuration 53
9.2. Component Weight Buildup 53
9.3. Component Positioning and Center of Gravity 54
9.4. The CG Envelope 57
9.5. Future Work in Configuration 58
10. Cost Analysis (MS) 60
10.1. Propulsion Cost 60
10.2. Air Frame Cost 60
10.3. Rated Aircraft Cost 60
10.4. Cost Analysis Results 60
11. Conclusion (AL) 62
12. References 63
iii
NOMENCLATURE
A Intake area
α Angle-of-attack
0α Angle-of-attack at zero lift
max,LCα Angle-of-attack for maximum lift
coefficient
AF Activity factor
AGL Above ground level
AR Aspect ratio
AUAS Agricultural unmanned aircraft
system
b Wing span
hb Span of horizontal stabilizer
vb Span of single vertical stabilizer
2,vb Span of dual vertical stabilizer
c Wing chord
c Mean aerodynamic chord of main
wing
DC Drag coefficient
iDC Induced drag coefficient
0DC Parasite drag coefficient
fC Skin-friction drag coefficient
CG Center of gravity
hc Chord of horizontal stabilizer
LC Lift coefficient
αlC Airfoil lift-slope
αLC Wing lift-slope
hLCα
Horizontal stabilizer lift-slope
max,lC Airfoil maximum lift coefficient
max,LC Wing maximum lift coefficient
l tailC Tail lift coefficient
TOLC Take off lift coefficient
fusmCα
Derivative of fuselage moment
with respect to angle-of-attack
pc Power coefficient
Tc Thrust coefficient
vc Chord of single vertical stabilizer
2,vc Chord of dual vertical stabilizer
z
C Normal force coefficient
z tailC Normal force coefficient of the
tail
maxzaC Max normal force coefficient
minzaC Minimum normal force coefficient
d Diameter of fuselage
D Propeller diameter
α
α
∂
∂ h Tail angle-of-attack derivative
α
α
∂
∂ p
Propeller angle-of-attack
derivative
zaC
α
∂
∂
Change in normal force
coefficient with respect to alpha
fδ Flap deflection angle
e Oswald’s span efficiency factor
FF Form factor
pF Propeller normal force
αpF Derivative of propeller normal
force with respect to alpha g Acceleration due to gravity
Fh Height of fuselage
hη Ratio of dynamic pressure at tail
to dynamic pressure at wing
pη Propeller efficiency
hp Horsepower
J Advance ratio
K Gust alleviation factor
pK Constant based on number
propeller blades
L Total lift force
λ Taper ratio
c25.0Λ Sweep angle at wing quarter
chord
tmaxΛ Sweep angle at maximum
thickness
DL / Lift-to-drag ratio
FL Length of fuselage
hl Distance of horizontal tail
quarter-chord behind center of
gravity
iv
vl Distance of vertical tail quarter-
chord behind center of gravity
M Moment
cgM Moment about aircraft center of
gravity
fusM Moment produced by fuselage
rootM Bending moment at wing root
wM Moment produced by wing
fwM δ Moment produced by wing
derivative with respect to flap
deflection angle µ Mass ratio
n Load factor (Structures)
Propeller rotation rate (Propulsion)
gn Gust load factor
q Dynamic pressure
RFP Request for proposal
majorr Large spray pattern turn radius
orrmin Small spray pattern turn radius
Q Interference factor
expS Exposed planform area
hS Planform area of horizontal
stabilizer
rootShear Shear force at wing root
refS Planform area of main wing
vS Planform area of vertical
stabilizer
wetS Wetted area
T Thrust
ct Airfoil thickness ratio
TOGW Take off gross weight
UAV Unmanned aerial vehicle
deU Derived gust velocity
V Velocity
cV Cruise Velocity
dV Dive Velocity
hV Horizontal tail volume coefficient
opV Operating velocity
tipV Propeller blade tip velocity
vV Vertical tail volume coefficient
Fw Width of fuselage
achX Longitudinal location of
horizontal stabilizer aerodynamic
center
acwX Longitudinal location of wing
aerodynamic center
acwX Location of wing aerodynamic
center non-dimensionalized by
wing chord
cgX Longitudinal location of aircraft
center of gravity
npX Location of neutral point non-
dimensionalized by wing chord
pX Location of propulsion unit
pX Location of propeller
aerodynamic center non-
dimensionalized by wing chord
tz Vertical distance between center
of gravity and thrust location
v
EXECUTIVE SUMMARY (AL)
The Agricultural Unmanned Aircraft System (AUAS) was proposed as an affordable
agricultural aircraft for use in crop-dusting. Conceptual Design yielded three configurations for
the AUAS: the conventional configuration, the canard with tractor configuration, and the canard
with pusher configuration. As outlined by the Request for Proposal, the AUAS must be capable
of taking off within 750 ft, spraying a field with dimensions of a half mile by 1000 ft, and then
landing. Because the aircraft must be low in cost and easy to operate and maintain, the main
design goals were set to minimize cost through reduction of weight and to maximize user
friendliness by designing a stable, simple plane.
After analysis, these configurations were sized to have takeoff gross weights of around
800 lb with spans of 20 ft and fuselage lengths ranging from 14-17 ft. The conventional
configuration was the lightest and shortest of the three. The two canard configurations were the
same weight, but the one with a pusher propulsion system was larger in dimensions. The
configurations were able to meet the RFP requirements, but to varying degrees. The
conventional was able to perform slightly better than the canard configurations in all the mission
segments due to its lighter body.
At this point, the leading design configuration is the conventional concept. However,
there are further improvements that can be made to better achieve the design goals, which
include designing a wing that produces more lift and reducing the weight and size of the plane.
These improvements can be attained through a more detailed design process, which could
develop the AUAS into a practical option for agricultural use.
1
1. INTRODUCTION (JN)
The Request for Proposal calls for the design of the Agricultural Unmanned Aircraft
System (AUAS), an airplane used for aerial application of liquid and solid particles on crops.[1]
Most existing agricultural aircraft are large, costly, and require complex supporting equipment.
Consequently, they are unsuitable for use in underdeveloped nations. There is therefore a need
for an inexpensive, easy to operate, and rugged crop duster for use worldwide.
The specific design constraints dictate that the design should be an unmanned, fixed wing
airplane, controlled by a pilot on the ground. The payload to be applied to a crop will consist of
100 liters of liquid chemical with density 64 pounds per cubic foot or 300 pounds of solid
particles with density 70 pounds per cubic foot. Each payload should be contained within a
hopper tank, which should be quick to load in the field. The equipment used to pump the liquid
and solid materials will be contained within a sphere with a one foot radius weighing thirty
pounds.
The operational and performance requirements dictate that the airplane should operate at
20 feet above the ground with reserve fuel for 20 minutes of flight after the mission is complete.
Additionally, the maximum landing and takeoff distance is 750 feet on a gravel runway with a
width of 50 feet. The stall speed should be the operating speed divided by 1.3. The airplane
should also be capable of one to two mile ferry flights at an altitude of 1000 feet.
The aircraft and all supporting equipment should be transported or towed by a standard
pickup truck. The design must allow future upgrades for greater endurance, more payload, and
higher altitude. It should be easy to operate, repair, maintain, and support. All costs associated
with its ownership and operation should be low. In addition, as many parts of the airplane as
2
possible, including the propulsion system, should be widely available. Finally, the airplane must
not endanger its operator or any surrounding people or property in the event of failure.
The mission profile for the design begins with a 5 minute warm-up and taxi. Next, the
airplane will take off and climb to 50 feet AGL. The airplane will then fly to the desired field
and descend to an altitude of 20 feet AGL. The cruise portion of the mission consists of spraying
a rectangular area measuring half a mile length by one thousand feet width with either the solid
or liquid payload. After the payload is dispersed, the airplane will align with the landing site and
land. This site can be either set on the other side of the field so that the plane has to be retrieved
later, or it can be the same site as takeoff, where the plane would have to climb back up to 50
feet, then align with the site and land. The mission profile is shown in Figure 1.1.
Figure 1.1. Mission profile with landing site on opposite side of the field
Given the mission profile and design requirements, Team Two-CAN determined several
design goals. These goals set the theme on which to base design decisions. The primary
attributes for which the design will be optimized are low cost and ruggedness. The final airplane
design will reflect the former of these ideals by minimizing the weight of the aircraft and by
using inexpensive parts and materials. The latter goal will be met by ensuring structural
soundness and ease of maintenance and repair.
3
2. CONFIGURATION SELECTION (KH)
2.1. Objectives
The goal of the configuration selection process was to identify three configurations that
were best suited to meet the requirements set forth in the RFP. At its most basic level, a
configuration consisted of a wing and a fuselage, to which was added empennage, a motor mount
and propulsion system, and landing gear. The configurations were down-selected based on the
following criteria derived from the design goals:
• The airplane must be easy to transport, hence fuselage length and half span should not
exceed 16 feet, provided that the wings are removed or folded for transport.
• Increasing the wingspan allows the airplane to cover a wider swath of field in a single
pass, meaning fewer passes are required. This reduces the range that must be flown.
• Purchase cost should be minimized. Because cost is closely tied to weight, weight should
also be minimized.
• The airplane should be easy to load and unload on the field. Hence, wing or empennage
configurations that obstruct large sections of the fuselage should be discouraged.
• The payload and other internal components should fit without an excess of extra space.
At the same time, provisions should be made for future upgrades including an increase in
payload capacity, so some extra space should be available.
2.2. Morphology
The fuselage/wing combinations, empennages, engine systems, and landing gear under
initial consideration are listed in Table 2.1. More obscure components such as gull or annular
wing, cruciform or H-tail, or mono-wheel landing gear were deemed noncompetitive and omitted.
4
Table 2.1. Morphological Chart Component Types
Monoplane Biplane Tandem Wing Wing/Body Joined Wing Flying Wing Blended Wing
Conventional T-Tail Canard Empennage Vertical Only None Dual Vertical
Landing Gear Tricycle Taildragger Bicycle + Skids
Propulsion Tractor Pusher Wing-Mounted
Each wing/body combination was matched with other components to produce the most
competitive and reasonable configurations. Seven configurations received further consideration:
• A conventional configuration, consisting of a mono-wing, conventional tail, taildragger,
and tractor.
• A tandem configuration, consisting of tandem wings, single vertical tail, tricycle gear,
and tractor.
• A biplane, consisting of biplane wings, conventional tail, tricycle gear, and tractor.
• A blended wing/body, consisting of the blended wing, conventional tail, tricycle gear,
and tractor.
• A joined wing configuration, consisting of joined wing and horizontal tail, a single
vertical tail, tricycle gear, and tractor.
• A flying wing configuration, consisting of a flying wing and single vertical tail, tricycle
gear, and tractor.
• A canard configuration, consisting of a monoplane, a canard and double vertical tails, a
pusher, and tricycle gear.
2.3. Configuration Selection
The seven configurations were then examined based on the criteria listed in Section 2.1.
It was noted that the wingspans on flying wings tend to be very large, often exceeding what
5
would be easily transportable. Also, the wing thickness generally is not sufficient to
accommodate the payload in a manner that is easy to load, and stability is difficult to achieve.
Thus, the flying wing configuration was eliminated. It was further noted that multi-wing
configurations such as tandem and biplane are useful for reducing the wingspan. In this case, the
wingspan would be reduced to a point where the swath width and range would be adversely
affected. Also, the multi-wing configurations left less of the fuselage easily accessible for
loading. Thus, both multi-wing configurations were eliminated. It was further noted that the
joined-wing configuration involved exceedingly complex geometry and thus would be difficult
to produce, maintain, and upgrade. Also, the majority of the joined-wing fuselage was
unavailable for loading. Thus, it was eliminated as well. Finally, it was noted that the blended
wing configuration was structurally more complex and heavier than the equivalent conventional
configuration, and the only advantage was slightly reduced drag and additional space that
probably wasn’t necessary. The blended wing configuration was therefore also eliminated.
Three configurations were derived from the remaining choices for detailed analysis.
The first configuration selected for analysis was the conventional mono-wing
configuration. It was selected not only based on its own strengths, but to serve as a point of
comparison for the other more complicated configurations. In this case, the wings would rotate
and swing backwards flush with the fuselage for ease of transport.
Because it was not immediately apparent how the engine position would affect the weight
and balance of a canard configuration, the remaining two configurations analyzed were canards,
one a canard with pusher and the other a canard with tractor. The canard with tractor consisted
of a mono-wing, canard, and single vertical tail. It was not immediately apparent which landing
gear configuration would be more feasible, so this decision was left for after weight and balance
6
analysis was completed. For transport, the wings would rotate and fold against the fuselage,
much like the conventional configuration. The canard with pusher was the most complicated
configuration under consideration. Because of likely tail/propeller interference, dual vertical
tails were used, and these were mounted to the wings via booms outside the propeller radius.
Also, in order to avoid the propeller passing through the spray stream, it was necessary to raise
the engine and propeller above the body axis. The vertical surfaces would need to be removed
for transport, and the wings would rotate and fold against the fuselage.
2.4. External Configuration
A rough fuselage size was also determined prior to initial sizing. A cylindrical fuselage
was selected to minimize drag, with a minimum diameter of 2 ft to allow the pump assembly to
fit. Most of the propulsion systems under consideration had a similar diameter. The fuselage
was to consist of three sections. The first was a motor mount section 3.5 ft in length with a slight
taper towards the front of a tractor or the rear of a pusher configuration. The second section was
a cylindrical payload compartment, the length of which would be determined in weight and
balance analysis. A diameter of 2.5 ft was selected to allow for structure around all major
components. The third section was a nose (pusher) or tail (tractor) cone which would taper from
the payload section diameter to a point. The length of the cone was dictated by the fact that a
deviation from the free stream of more than 12 degrees would produce flow separation, as
specified in Raymer.[2]
Given the payload section diameter, a length of 6.0 ft was deemed
sufficient. Figures 2.1-2.3 show the configurations under consideration.
7
Figure 2.1. Three-view of the conventional configuration.
8
Figure 2.2. Three-view of the canard with tractor configuration.
9
Figure 2.3. Three-view of canard with pusher configuration.
10
3. SIZING ANALYSIS (AK, CC)
3.1. Initial Sizing
Initial sizing analysis was completed for three aircraft configurations; a conventional
tractor, canard tractor, and canard pusher. This analysis was done by first creating a conceptual
aircraft. The specifications can be found in Table 3.1.
Table 3.1. Characteristics of Conceptual Design Aircraft
Wing Span
(ft) Chord (ft)
Cruise
Velocity
(mph)
Loiter
Velocity
(mph)
Swet/Sref L/Dmax
20 3 65 60 5 10.5
Using the initial sizing procedures outlined by Raymer, the individual segments of the
mission were evaluated to determine the weight fractions of the mission, Wi/Wi-1. The takeoff
and landing weight fractions were taken from empirical data, however the climb value was
modified. Since the climb was only 50 ft, it was not very logical that the empirical value of
0.985 would hold true since the climb would last only a few seconds. Therefore, an adjusted
value of 0.998 was adopted. The cruise weight fraction was calculated using the equations and
methods discussed in initial sizing section of Raymer.[2]
After calculating the individual weight fractions as shown above in Table 3.2, their
product was taken to find the total weight fraction over the mission. This value was used to find
the fuel weight fraction as shown in Equation 3.4 where 1.06 is used to account for the trapped
fuel. This fuel weight fraction was then put into Equation 3.5. From here, an iteration was done
by guessing an initial weight, evaluating Equation 3.6 at that value, evaluating Equation 3.5 at
that We/Wo. Equation 3.6 was evaluated based on empty weight fractions of other unmanned
aerial vehicles; this provided a more reasonable answer than other coefficient sets for manned
aircraft.[3]
11
Table 3.2. Mission Segment Weight Fractions
Range
(mi)
Endurance
(min)
Velocity
(mph) L/D
SFC
(1/hr) Wi/Wi-1 Wi/Wo
0 Warm-up -- -- -- -- -- 1.0 1.0
1 Takeoff -- -- -- -- -- 0.97 0.97
2 Climb/Descent -- -- -- -- -- 0.998 0.968
3 Cruise/Turn 25 -- 65 10.5 2.41e-5 0.996 0.964
4 Climb -- -- -- -- -- 0.998 0.962
5 Loiter -- 20 60 9.1 3.17e-5 0.996 0.958
6 Landing -- -- -- -- -- 0.995 0.954
6
0
1.06 1f
o
W W
W W
= −
[3.4]
0.07950.916eo
o
WW
W
−= [3.5]
o
o
f
o
o
e
equippayloado WW
WW
W
WWWW +++= [3.6]
This iteration led to a TOGW of approximately 800 lb for the conventional tractor
configuration. Initial sizing analysis was also conducted on he canard pusher and canard tractor
configurations, but the conceptual designs were nearly identical, and since initial sizing is an
extremely generalized process that takes few configuration specifics into consideration, their
results were identical. Therefore, those tabulations and calculations have been omitted. The
final TOGW of all three configurations was around 800 lb. The weight breakdown is shown
below in Table 3.3. It is worth noting that all three configurations have identical results since
canards and conventional aircraft are extremely similar.
Table 3.3. Pertinent Weights based on the Initial Sizing of the Conceptual Designs
Configuration TOGW (lb) Empty Weight
(lb)
Empty Weight
Fraction
Mission Fuel
(lb)
Fuel-Weight
Fraction
all 800 432 0.54 39.2 0.049
3.2. Constraint Analysis
Upon completion of the initial sizing which generated a TOGW of approximately 800 lb,
the next step of conceptual design was to perform a constraint analysis which ultimately
12
produced a design point for a practical wing and power loading given the mission performance
constraints outlined by the RFP. The constraint analysis was performed on the four vital
segments of the mission profile: the take-off, cruise, sustained turn, and landing.
The results of the constraint analysis can be found in Figure 3.1, which includes the
design point selected. The design point chosen lies within the allotted design space, however, it
does not fall at the absolute minimum wing loading permitted because the RFP places a physical
size constraint by stating that the aircraft must be transported using a mid-size pick-up. Given
this additional constraint, it is increasingly difficult to design an aircraft with wings that have a
large enough reference area to achieve a low wing loading while maintaining ease of
transportation and operation which would be severely hindered if an operator was required set-up
long or bulky wings before flight. This design point, like the initial sizing, was found by
giving little consideration to the specific parameters unique to the different configurations.
(11.5, 0.1)
0
0.02
0.04
0.06
0.08
0.1
0.12
0.14
0.16
0.18
0.2
0 5 10 15 20 25
(W/S)o (lbf/ft2)
(P/W)o
Take-off Constraint
Cruise Constraint
Sustained Turn Constraint
Landing Constraint
Design Point
Figure 3.1. Constraint analysis diagram.
13
4. PERFORMANCE (CC)
4.1. Introduction to Performance
The primary objective for the AUAS is to dispense a payload of solid or liquid over a
field measuring one half mile by 1000 ft. Given this governing stipulation, it was necessary to
devise a strategy to accomplish said goal in an effective, efficient, and simple manner. This
section describes both the qualitative and quantitative analysis used to devise a spray pattern that
is viable to meet the constraints set by the RFP.
Additional constraints of the RFP provide that the UAV must be able to perform the
mission profile outlined in Section 1. The speed at which the UAV must operate should be 1.3
times the stall speed and the UAV must also be able to satisfy the requirement of performing
short ferry flights between 1 and 2 miles without payload at 1000 ft AGL.
Given this mission profile, it should be noted that the most essential flight segments are
the take-off and the cruise in which the UAV is dropping its payload and employing a tactical
flight path with multiple turns. The climb segments are thus omitted from the conceptual design
process as they are not severely constrained and quite insignificant in relation to the overall
demands. Consideration will, however, be given to the loiter requirement of 20 minutes of
reserve fuel. Additionally inclusive in this study is the assessment of the variations in
performance capabilities between the different proposed configurations.
4.2. Takeoff Analysis
As set by the RFP, the vehicle must be capable of a 750 ft takeoff distance. This
requirement is open-ended in that it does not imply the necessity to design for obstacle clearance
ability beyond the runway confinement. Therefore, for the purpose of the conceptual design, it is
assumed that 750 ft is the maximum allotted ground roll distance.
14
The determination of the take-off distance is appropriately acquired using a force body
diagram of the plane at horizontal with respect to ground. The acceleration of the plane is given
by subtracting the drag and friction forces from thrust and dividing by the UAV’s mass. The
complication of this calculation lies in determining these force values since they are all functions
of the velocity which is increasing at take-off. The thrust was therefore modeled by assuming
the ideal case that it is equal to the power multiplied by propeller efficiency and then divided by
the velocity. In this assumption, the propeller efficiency was held constant, and power was
assumed to be at maximum of 70 hp, a value deduced from the constraint analysis. The friction
force was modeled by assuming a friction coefficient given the known runway composition, then
predicting normal force for the maximum take-off weight, found from initial sizing, then
subtracting the lift. The lift, as well as the drag, is strongly dependent on velocity and wing
geometry. This gives rise to yet another unknown variable, the lift coefficient in the take-off
configuration. However, given that all forces have now been represented as functions of velocity
and the unknown lift coefficient at take-off configuration, a MATLAB™
program was created to
determine the minimum lift coefficient needed to get the plane off the ground within the runway
constraint. The integral of velocity divided by acceleration with respect to velocity is the ground
roll distance, and by setting this distance equal to a maximum of 750 ft and numerically
integrating from 0 feet per second to VTO. This process was performed over a range of CL,TO
values which directly govern the necessary take-off velocity at which lift equals drag. The
minimum take-off lift coefficient was determined when the ground roll plus rotation distance
became equal to or just less than 750 ft. This approach required the estimation of CDo from
Aerodynamics. The calculated minimum CL,TO values and maximum take off velocity for the
different configurations can be seen in Table 4.1 for comparison. The differences arise from the
15
different proposed CDo and TOGW values for each configuration where TOGW is assumed to be